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3.15.1 Airframe strength and elasticity Aircraft structures are designed to be as light as possible with some degree of structural flexibility while still providing adequate strength for the aircraft's operational category. To receive type certification, the design of a general aviation or recreational aircraft must conform with certain standards — among which are the in-flight structural load minimums — for the category in which the aircraft may be operated. In FAR Part 23, a recognised world standard for light aircraft certification, the minimum load factors that an aircraft at maximum take-off weight [MTOW] must be designed to withstand are: +3.8g to –1.5g (or –1.9g) for the normal operational category (which would include most factory-built recreational aircraft). +4.4g to –1.8g (or –2.2g) for the utility category (which includes most GA, and perhaps some RA, training aircraft). +6g to –3g for the acrobatic (i.e. aerobatic) category. +4g to –2g for the Light Sport Aircraft category manufactured in compliance with ASTM F2245-07 standard specification for design and performance of a Light Sport Airplane. Sailplanes and powered sailplanes are generally certificated in the utility or acrobatic categories of the European Joint Airworthiness Requirements JAR-22, which is the world standard for sailplanes; aerobatic sailplanes have limit loads of +7g and -5g. For more information see 'Limiting loads and ultimate loads'. There is an increasing risk of failure when exceeding the minimum load factors, and each instance of excessive loading will compound the failure risk. We use load factors in terms of g for convenience, but what we are also considering is total aerodynamic loading — remember that dynamic pressure increases with the square of the velocity; i.e. dynamic pressure = ½rV². Notes: 1. Uncertificated minimum ultralight aircraft, even with their low wing loading of perhaps 12 kg/m², can be overstressed readily just by flying at maximum level speed and increasing g in a pull-up (positive g) or a push-over (negative g). 2. Many aircraft are type certificated in both normal and utility category, and some are certificated in those plus the acrobatic category. In this case, the MTOW and cg limits are not fixed values, but vary according to the flight operating category. See the table in weight/cg position limitations. Weight and balance There are fixed limits to the payload an individual aircraft may carry safely. The payload must be distributed so that the aircraft's balance — the position of the aircraft's centre of gravity — is maintained within calculated limits. In addition, there is a maximum safe operating weight permitted by the aircraft designer. However, for many recreational aircraft, the MTOW will be limited by national legislation, which has nothing to do with aeronautical engineering. The aircraft's weight and balance very much affect control and stability at high speeds. Excess weight reduces the designed structural load limits, while cg positions outside the designated fore and aft limits may enhance unfavourable reactions to aerodynamic loads, affect stability, reduce controllability, or delay (or prevent) recovery from unusual or high-speed situations. Aeroelastic effects Elasticity is a property of certain materials that enables them to return to their original dimensions after an applied stress has been removed, see the notes on 'stress and strain' in the 'Builders guide to safe aircraft materials'. Elastic structures have a natural frequency of vibration and all aircraft structures exhibit some degree of elasticity; that is, they distort or deform a little, changing shape — flexing, elongating, compressing, bending and/or twisting — under applied aerodynamic loads; each type of distortion produces a particular mode of vibration. Transient structural distortions also contribute to a change in the aerodynamic forces, so the distortions and forces are mutually dependent. This is particularly so with the wings and tailplane. Wings have a low frequency bending mode of vibration where the tips flex up and down (i.e. flap) relative to the wing root, under changing flight loads — in turbulence for example. The degree of oscillation or flapping is more pronounced with high aspect ratio wings. While bending upward the wing adds a vertical velocity* to its forward velocity — the true airspeed — which results in a decreasing angle of attack (aoa) reducing the lift of the up-moving portion of the wing and thus causing an aerodynamic damping of the flapping oscillation. Similarly a downward bending motion results in an increasing aoa, increasing the lift of the down-moving wing and again causing an aerodynamic damping of the flapping oscillation. *A similar resultant velocity concept to a vertical gust encounter. Wings also exhibit a higher frequency torsional mode of vibration where they twist about the wing's elastic axis as the centre of pressure moves chordwise and consequently produces a spanwise variation in the aoa and changes the lift force and its distribution. If the centre of pressure moves forward it can then again increase the wing twist, aoa and lift, developing a non-stable situation. Twisting and bending distortions result in independent oscillations or vibrations and alter the effectiveness of lifting surfaces, though structural and inertial forces provide a natural positive damping that normally keeps vibratory energy in check. The elastic axis is defined as the line along the span of the wing where no torsion occurs when a loading is applied to the wing. A lifting force centred aft of the elastic axis will tend to twist the outer wing leading edge down, reducing aoa and thus aerodynamic force. A lifting force centred forward of the elastic axis will tend to twist the leading edge up, increasing the aoa and the aerodynamic force. The wing aerodynamic centre is usually designed to be close to or behind the elastic axis; if the aerodynamic centre is forward of the elastic axis then wing twist will increase aoa leading to the non-stable situation described in the preceding paragraph. The degree of torsional distortion is dependent on (1) the area of wing surface affected, (2) the distance between the aerodynamic centre and the elastic axis plus (3) the torsional stiffness (rigidity) of the surface. The torsional stiffness designed into the wing resists twisting, and structures usually revert to the normal status when the load is normalised. Aeroelasticity may lead to some problems at high speed, but reducing elasticity means increasing rigidity, which perhaps involves an unwarranted increase in structural weight. So, aircraft structural engineering must be a compromise between rigidity and elasticity. 3.15.2 Aerodynamic reactions to flight at excessive speed Flutter Wing structures are akin to a very-low-frequency tuning fork extending from the fuselage. When a tuning fork is tapped, the fork vibrates at a particular frequency; the stiffer the structure, the higher its natural frequency. The natural frequency of a wing or tailplane structure may apply another limiting airspeed to flight operations related to a self-exciting interaction between elastic, aerodynamic and inertia forces that result in 'flutter' of control surfaces and the structure to which the surface is attached. For example, when the airflow around a wing, tailplane or control surface is disturbed (by aerodynamic reactions, turbulence or pilot inputs) the structure's elastic reactions – twisting and bending – may combine as an oscillation or vibration of the structure that will quickly damp itself out at normal cruise speeds because of the structure's resistance. It is possible that the oscillation does not damp out but is sustained at a constant amplitude (perhaps felt in the airframe as a low-frequency buzz) that is not, in itself, dangerous but may contribute to structural fatigue. At some higher airspeed — the critical flutter speed, where the oscillations are in phase with the natural frequency of the structure — the oscillations will not damp out but will become resonant, rapidly increasing in amplitude. (Pushing a child on a swing is an example of phase relationships and amplification.) This flight resonance – flutter – is very dangerous, and unless airspeed is very quickly reduced, the increasing aerodynamic forces will cause control surface (or even wing) separation within a very few seconds. In 1966 NASA recorded a 24-second video of an in-flight flutter test on a Piper PA30 Twin Comanche demonstrating how rapidly stabilator oscillations increase in amplitude; Google 'NASA flutter video'. The following is an extract from an article by William P. Rodden which appeared in the McGraw-Hill Dictionary of Science and Technology; it provides a succinct description of flutter: "Flutter (aeronautics) — An aeroelastic self-excited vibration with a sustained or divergent amplitude, which occurs when a structure is placed in a flow of sufficiently high velocity. Flutter is an instability that can be extremely violent. At low speeds, in the presence of an airstream, the vibration modes of an aircraft are stable; that is, if the aircraft is disturbed, the ensuing motion will be damped. At higher speeds, the effect of the airstream is to couple two or more vibration modes [e.g. bending plus twisting ... JB] such that the vibrating structure will extract energy from the airstream [JB's emphasis]. The coupled vibration modes will remain stable as long as the extracted energy is dissipated by the internal damping or friction of the structure. However a critical speed is reached when the extracted energy equals the amount of energy that the structure is capable of dissipating, and a neutrally stable vibration will persist. This is called the flutter speed. At a higher speed, the vibration amplitude will diverge, and a structural failure will result." So, flutter is a vibrational instability that (if the structure is not sufficiently stiff) is generally related to the aerodynamic forces and thus the airspeed; but there are many flutter modes. Providing high torsional stiffness in an airframe structure — particularly a high aspect ratio wing — may incur weight penalties that are unacceptable for those aircraft whose MTOW is limited by national legislation, rather than normal design parameters. Mass inertia and the status of the control actuation systems are also involved in flutter development. Consequently ailerons, elevator or stabilator and the rudder (in that order) should be considered for mass-balancing, i.e. their centre of gravity is made coincident with their hinge centre line. This limits the mass moment of inertia so that the control surface does not rotate about its hinge line when the main structure moves; e.g. when the wing bends upward the aileron does not rotate downward but maintains the same floating position relative to the wing. It may be acceptable for the control surface to be over-balanced; i.e. its cg is slightly forward of the hinge line but under-balancing may achieve little. Mass-balancing of the control surfaces, including the rudder, should prevent flutter of that control surface, but the possibility of, for example, wing flexing/twisting flutter might still exist. Mass-balancing of ailerons might be accomplished by attaching moulded lead weight/s, or a lead-filled steel tube, within the nose of the control structure forward of the hinge centre line. As the moment arm, between that centre line and the centroid of the added weight, could be quite short the balance mass needed could be twice the mass of the unweighted control, so it is possible the end effect could be close to tripling the total weight of the aileron. Mass balancing of elevators might be achieved using a weight on an arm contained within the tail structure or within control horns, while a rudder might also be balanced by a weight within a control horn. The friction within aircraft control surface actuating systems adds to the damping of control surface oscillations and this damping ability increases as the oscillation frequency increases. So, it is important that all parts of the control actuating systems are made as rigid and secure as possible and checked to ensure that rigidity is always maintained so that control surfaces cannot deflect without a corresponding movement in the cockpit control. The possibility of destructive flutter increases if any of the following conditions exist: wear in control surface hinges, pulleys, fairleads or guides lack of tension, wear or slop in actuating pushrods/cables/conduited push-pull cables/cranks/torque tubes/turnbuckles safetying wire improperly installed faulty trim tabs. Also water or ice inside control surfaces or absorbed within a foam core; mud outside; additional surface coatings applied after mass balancing; tail buffeting caused by unsteady airflow related to, for example, alterations to the engine exhaust system; or other system anomalies that alter structural reactions also play a role in flutter development. Also see AC43.13-1B Chapter 7 'Aircraft hardware, control cables, and turnbuckles'. This is an extract from an RA-Aus accident investigation report: "(Witnesses) observed the aircraft in a steep dive at what appeared to be full power. The port wing appeared to detach from the aircraft ... The wing that tore away from the fuselage had the attach points intact but had pulled the mountings out of the top of the cockpit. This action would have released the door, which landed close to the wing. The wings were intact but the ailerons were detached. There was no delamination of the fibreglass structure. The ailerons were not mass balanced. The (prototype) aircraft was a conventional design being a high wing, monoplane of composite construction. While the fuselage was a proven design the pilot /builder had designed his own wing including the aerofoil section. The workmanship was excellent and there is no evidence of any lack of structural integrity. The eyewitnesses reported seeing a sort of 'shimmying' from the aircraft. It is believed that this shimmying was aileron flutter which led to the detaching of both ailerons. This same flutter condition would account for the massive forces required to detach the wing from the aircraft in the manner that occurred. Flutter could have been triggered by the wing aerofoil design combined with the manoeuvre the pilot was conducting or from the aileron control design ... The aircraft suffered a massive inflight structural failure almost certainly caused by severe aileron flutter and the aircraft speed in the dive. Any flutter would have been exacerbated by the lack of mass balancing." Vne — the standard limiting airspeed If an aircraft is operated within its specified flight envelope, observing the limiting accelerations and control movements, and maintaining airspeed commensurate with atmospheric conditions, then the only possibilities of in-flight structural failure relate to: improper modification, repair or repainting of the structure control actuating system deficiencies cumulative strain, or minor damages, in ageing aircraft failure to comply with the requirements of airworthiness notices and directives poor care and maintenance of the airframe. Flight at airspeeds outside the envelope (or at inappropriate speeds in turbulent conditions, or when applying inappropriate control loads at high speed) is high risk and can lead to airframe failure. Vne is the IAS, specified by the designer, which should never be intentionally exceeded in a descent or other manoeuvre. For a fuller description of Vne and how it is calculated see 'How fast is too fast?' in the 'Decreasing your exposure to risk' tutorial. Wing divergence Wing divergence refers to a state where — at very low angles of attack and high speed (when the nose-down pitching moment is already very high) — pressure centres develop, which push the front portion of the wing downward and the rear portion upward. This aerodynamic twisting action on the wing structure — while the rest of the aircraft is following the flight path — further decreases the aoa and compounds the problem. The action finally exceeds the capability of the wing/strut structure to resist the torsional stress, and causes the wing to separate from the airframe with no warning. This could be induced if turbulence is encountered at high speed. Control reversal As airspeed increases, control surfaces become increasingly more effective. They reach a limiting airspeed where the aerodynamic force generated by the ailerons, for example, may be sufficient to twist the wing itself. At best, this results in control nullification; at worst, it results in control reversal. For example, if the pilot initiates a roll to the left, the downgoing right aileron will twist the right wing, reducing its aoa and resulting in loss of lift and a roll to the right, probably with asymmetric structural loads. All of which would make life difficult when attempting to roll the wings level during recovery from a high-speed dive. Many of the uncertified minimum ultralights, and perhaps some of the certificated aircraft, have low torsional wing rigidity. This will not only make the ailerons increasingly ineffective with speed (and prone to flutter), but will also place very low limits on Vne and g loads. Vne may be so low that it can be achieved readily in a shallow descent at 75% power. Effect of wing washout Wings incorporating geometric washout have a significantly lower aoa towards the wing tips. At high speed when the wing is flying at low aoa, there are high aerodynamic loads over the wings. However, the outer sections could well be flying at a negative aoa and the reversed load in that area will bend the wingtips down, possibly leading to outer spar fracture. See the accident technical report below. Vertical gust shear and gust loads The effective aoa of an aircraft encountering an atmospheric gust with a significant vertical component (updrafts, thermals, downdrafts, microbursts, macrobursts and lee waves) will be increased momentarily if the air movement is upward relative to the aircraft's flight path, or decreased momentarily if the air movement is downward. Thus, an updraft will increase CL and lift, increasing the aerodynamic loading and lead to an upwards acceleration of the aircraft. The magnitude of the acceleration is determined largely by the change in aoa, the aircraft speed (the higher the speed, the greater is the g load), the design wing loading and the aspect ratio. The lower the design wing loading and/or the higher the aspect ratio, the greater is the change in load factor for a given increase in aoa and the easier it is to overstress the wings at high speed. The effects of shear and gust loads are expanded in the section on wind shear and turbulence. Other effects It is not just the preceding items that may be a problem at high speed. The maximum speed may be limited by the ability of the fuselage to withstand the bending moments caused by the loads on the tailplane necessary to counter the wing's substantial nose-down pitching moment at very low aoa, or the aoa changes due to vertical gust shear, or the extreme loads caused by a high speed pull-up. Applying rudder in a high speed pull-up applies twisting loads to the rear fuselage. Even a very small bird can cause severe damage in a high-speed bird-strike. When nearing the zero-lift angle of attack in a high-speed descent, many cambered wings suddenly experience a strong nose-down pitching moment and the aircraft will 'tuck under' rapidly; this will certainly make the pilot wish she/he was somewhere else. The symmetrical aerofoil wings often used in aerobatic aircraft don't have this problem. Also, the possibility of a runaway propeller in a high-speed dive is always there for those aircraft with a constant-speed propeller governor or perhaps an in-flight adjustable system. The following is a condensed version of an Australian Transport Safety Bureau Technical Analysis Occurrence Report. Note: the Coroner's findings in relation to the fatal accident near Atherton does not support any view that the accident was caused by pilot mishandling; rather, the Coroner's "preference is towards port side wing tip separation as a consequence of the un-airworthy state of the aircraft ..." "An Airborne Edge microlight aircraft impacted terrain during a 2005 flight to Atherton, in Far North Queensland. The pilot, the sole occupant of the aircraft, was fatally injured. In 2006 a similar Airborne Edge aircraft impacted terrain at Cessnock, New South Wales, also fatally injuring the pilot, the sole occupant of the aircraft. In both instances, RA-Aus initiated safety investigations to determine contributing factors to these accidents. During the course of these investigations, similarities in the structural failures of both aircraft were observed. In addition, a third accident involving an Airborne aircraft registered with HGFA with similar structural failure was identified. This accident had occurred in 1996 in Hexham, NSW. In order to determine possible connections between all three accidents, ATSB was asked to conduct technical examination and analysis on recovered parts from the Atherton and Cessnock accidents, to assist the RA-Aus investigation. Information regarding the 1996 accident was taken from coronial findings. In all three accidents, the failure of the main wingspars had occurred near the wingtip. Qualitative analysis of the structural design and loading of the part during this safety investigation and the examination of the coronial findings from the Hexham accident, revealed that all main wingspars had failed under negative G loading. Such loading was likely if the aircraft entered or encountered flight conditions outside the manufacturer's specified flight envelope. Examination of material characteristics of the failed wingspars did not show evidence of material deficiencies that could have contributed to these accidents. The manufacturer's operating handbook prohibited all aerobatic manoeuvres including whipstalls, stalled spiral descents and negative G manoeuvres. The manual specified that the nose of the aircraft should not be pitched up or down more than 45 degrees, that the front support tube of the microlight and the pilot's chest limit the fore and aft movement of the control bar, and that the aircraft should not exceed a bank angle of 60 degrees. Review of photographs of the Airborne Edge, indicate that the wing adopts a degree of twist while in flight. Twist will effect the load distribution by shifting some of the lift from the tips inboard (i.e. more lift is generated in the middle of the wing). Given the structural restraint of the tip struts and battens located at the tip of the trailing edge of the wing, the aerofoil at the wing tip must adjust and try to align with the relative airflow. This results in a smaller amount of lift generated near the wing tips due to a reduced angle of attack to the relative airflow." (Or an aoa reduced below the zero lift aoa, i.e. reversed lift ... JB) 3.15.3 Recovery from flight at excessive speed Generally, excessive speed can only build up in a dive, although just a shallow dive can build speed — and rate of descent — quite quickly. The table below is a calculation of the rate of descent after a few seconds at dive angles of 10°, 30° and 45° for a moderately slippery light aircraft. Dive angle Airspeed (knots) Rate of descent (fpm) 10° 100 1700 30° 150 7500 45° 180 12 500 Recovery from an inadvertent venture into the realm of flight near, or even beyond, Vne is quite straight-forward, but requires pilot thought and restraint in initiating recovery procedures, particularly so if the aircraft is turning whilst diving. Considerable height loss will occur during recovery, so the restraint is required when terra firma is rapidly expanding in the windscreen. Halt the buildup in airspeed by closing the throttle. Unload the wings to some extent by moving the control column to the neutral position or just aft of it. Keep the slip ball and the ailerons centred — the twisting action of excess rudder at very high airspeed may strain the tailplane and rear fuselage. Gently roll off any bank while using coordinated rudder; this will ensure the total lift vector is roughly vertically aligned. Maintain the control column position at neutral or slightly aft to avoid any asymmetric loading arising from simultaneous application of aileron and elevator at high speed. When the wings are level, start easing back on the control column until you are pulling the maximum load factor for the aircraft : +3.8g or +4.4g, perhaps less for some ultralights. Do not pull back so harshly that the aircraft enters a high-speed stall. Hold the applied loading near the maximum until the aircraft's nose nears the horizon, then level off. The aircraft will have sufficient momentum to reach this position before opening the throttle. If you have ample height at the commencement of recovery, then there is no need to pull such high g — particularly if the atmosphere is bumpy when gust loads, added to the high manoeuvring g, may prove excessive. In aircraft not certified for aerobatics, it is best to wait until airspeed is less than Va before pulling g — if circumstances permit. A problem with this procedure is that most light aircraft do not have an accelerometer [g-meter] fitted, so it is difficult to judge the g being pulled. However, if properly executed 60° steep turns are practised, then some idea of the 2g load on your own physiology can be gained. At the higher end of acceleration the average fit person will probably start feeling the symptoms of greyout by 4g. 3.15.4 Recovery from a spiral dive In a well-developed spiral dive, the lift being generated by the wings (and thus the aerodynamic loading) to provide the centripetal force for the high-speed diving turn, is very high, and much of it is directed inward. The aircraft is near the extremes of its design flight envelope, with very high aerodynamic loading and very high speed, well above Va. The pilot must be very careful in the recovery from such a dive, or damaging structural loads will be imposed. If rearward stick force is applied to pull the nose up while the aircraft is turning, the result will be a tightening of the turn and further lowering of the nose, thus dramatically increasing the applied loading or possibly prompting a very punishing high-speed stall. Also aileron to level the wings must be applied with restraint, the aileron on the lower wing will increase the aerodynamic force on the portion of wing ahead of it and move the centre of force further towards that wingtip, so increasing the moment of force at the spar root. The downgoing aileron is also applying a twisting force to the outer wing structure. Sudden or excessive aileron deflection at airspeeds well above Va could well lead to outer wing or full wing separation. Control reversal could also be a factor, see 'High-speed control reversal: will it always roll in the direction you want?' The recommended procedure — for a fixed undercarriage aircraft without propeller pitch control — is: Reduce power. Carefully centralise controls: the forward movement of the control column will partially unload the wings. Smoothly level the wings with aileron while the rudder and elevators are held in the neutral position. As the wings become level with the aircraft still diving at high speed, much of the lift that was providing the centripetal force will now be directed vertically (relative to the horizon); and if up elevator is applied, the aircraft may start a high g pitch-up — even into a half loop. Thus to prevent this, the pilot must hold the elevators in the neutral position while rolling level or even applying further FORWARD stick pressure — before applying aileron — to reduce aoa; but not below the zero-lift aoa, i.e. the load factor must remain positive. At high speed, the stick force required will be high, but the position of the elevator trim should not be altered. Also it is probably not wise to apply two controls simultaneously at very high speeds because of the consequent asymmetric airframe loading. Read this Australian Transport Safety Authority analysis of an inflight breakup most likely caused by excessive control force during spiral dive recovery. The theme common to all problems encountered when moving at very high speed is that there is no warning and little time to do anything about it! The only safe procedure is not to push the high-speed end of the envelope at any height: make gentle, smooth control movements and avoid asymmetric flight loads and never put yourself in the position where you may encounter non-visual flight conditions at low levels. 3.15.5 Notes: compressibility of airflow and Mach number These notes have little value for the recreational aviator, but are included for interest. Except for a slight EAS correction to IAS/CAS, and the possible propeller effects, the compressibility/elasticity of airflow (i.e. the density change resulting from pressure disturbances) does not have any significant airframe aerodynamic effects for aircraft operating at speeds below 200 knots TAS and altitudes below 10 000 feet. Pressure disturbances, or waves, propagate through the atmosphere in all directions, at the speed of sound. Mach 1.0 is the notation for the speed of sound. For aerodynamic purposes airflow speeds are classified within five ranges: Hypersonic flow — airflows greater than Mach 5.0 Supersonic flow — airflows between Mach 1.5 and Mach 5.0 Transonic flow — airflows between Mach 0.8 and Mach 1.5 Subsonic flow — airflows between Mach 0.3 and Mach 0.8 Incompressible flow — airflows below Mach 0.3 The term 'incompressible flow' doesn't mean that air is incompressible; it just indicates that at flow speeds below Mach 0.3 (30% of the speed of sound or about 200 knots TAS), local density variations within the flow — due to compressibility — are insignificant; so aerodynamicists can assume constant density within the flow. At subsonic velocities, significant density changes may occur in the airflow around wings, which will produce flow separation and a turbulent wake — wave drag. The associated drag coefficient builds rapidly at airspeeds above Mach 0.75 then reduces as Mach 1.0 is exceeded. The speed of sound in the atmosphere varies with air temperature. The Mach number is the measure of an aircraft's TAS in relation to the ambient speed of sound. For example, Mach 0.6 indicates that the aircraft's true airspeed is 60% of the speed of sound. The speed of sound is proportional to the square root of the absolute temperature. In the ISA, Mach 1.0 at sea level = 663 knots, and temperature at sea level = 15 °C [288 K]. Thus, if the temperature = −36 °C (237 K) then the ambient Mach 1.0 = 663 × √237/√288 = 601 knots. Thus, Mach 0.60 at 15 °C would be 398 knots TAS, while Mach 0.60 at −36 °C would decrease to 360 knots TAS. Below the tropopause — the speed of sound decreases as altitude increases. A machmeter is an instrument that measures and compares the speed of the aircraft and the speed of sound, using the outside air temperature. It adjusts for actual air density but is still subject to the same position errors as the ASI. The machmeter is usually incorporated within an ASI; the numeric Mach appears in a small window within the ASI dial. You may see references to design diving speed presented as 'Vd/Md' which indicates the speed may be expressed as IAS or Mach number. Other reference airspeeds are presented in similar fashion. For interest, the following table is the maximum permissable speed/altitude for a late 1940s/early 1950s piston-engined naval fighter — the Seafire 47: Altitude feet Max. IAS knots Mach no. Approx. TAS Sea level – 10 000 455 0.78 505 10 000 – 15 000 410 0.78 495 15 000 – 20 000 375 0.78 485 20 000 – 25 000 340 0.78 472 25 000 – 30 000 300 0.78 459 30 000 – 35 000 270 0.78 448 35 000 + 240 0.78 432 Subsonic jet transport aircraft are designed to cruise close to their maximum allowable speed — Vmo/Mmo. Vmo is the limiting indicated airspeed and Mmo is the limiting Mach number. Mmo is probably between Mach 0.80 and Mach 0.85. In normal operations the limiting airspeed is Vmo, up to a change-over pressure altitude (perhaps around 25 000 feet). Above this altitude Mmo becomes the limiting speed value because of compressibility problem restraints. Vmo could be shown as a fixed red line on the ASI (or 'Mach/Airspeed Indicator') but, because the speed of sound decreases as altitude increases, Mmo can't be represented by a fixed marking on the indicator. So, a moving red-and-white striped pointer, the 'barber pole', shows the limiting Vmo/Mmo varying with altitude. It shows the IAS corresponding to the lower of Vmo or Mmo for the current altitude. For further explanation read this Boeing flight operations review document. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.14.1 The landing sequence In this module we will look at the common factors to be considered in landing a normally configured, three-axis, fixed-undercarriage, nosewheel or tailwheel aircraft, which may or may not be flap-equipped. Aircraft designed with full 'short take-off and landing' [STOL] capability will use slightly different techniques in some parts of the approach and landing. There are differing landing procedures or techniques, or combinations thereof, applicable to airfield dimensions and surface conditions: normal landing short-field landing soft-field landing. The basic landing sequence is varied, according to prevailing conditions (and there is a varying degree of alignment correction to allow for the crosswind component of the wind velocity), but it usually has four parts: Joining the circuit pattern of the airfield, during which the aircraft is decelerated from cruise speed to circuit speed, the airfield is visually checked for serviceability and obstructions, surface wind direction ascertained from observation of the windsock(s), the whereabouts of other traffic is established, the landing direction and approach is planned and the pre-landing cockpit checks are carried out in a logical sequence. The approach to the landing, during which the aircraft is decelerated from circuit speed to the reference indicated approach speed [Vref], configured for landing, then finally stabilised at a constant speed and rate of descent with wings level and aligned — so that the flight path traced over the ground, during the final approach, is on the same line as the intended ground roll-out path. The stabilised approach should be established before the aircraft is at a height 300–400 feet above the runway/airstrip/landing area. Once established, only slight movements of the flight and engine controls should be necessary to maintain the approach. The flight path passes over an imaginary 50 feet high screen, placed at a short distance before the airstrip threshold. A transition period, where both the rate of descent and the forward speed are slowed during a 'round-out' or 'flare' prior to touchdown. The touchdown and subsequent ground roll, after which the aircraft is turned off the landing area at an appropriate taxiing speed. The arrival is complete when the aircraft is properly parked, the engine is properly shut down, any passenger is safely disembarked and the aircraft is secured. The most favourable conditions for optimum landing performance at, or near, maximum weight are: a pilot who exercises sound judgement, and follows the rules and recommended procedures a surface of ample length, which is dry and level, or with a slight upslope a low density altitude; i.e. low elevation and low temperature a smooth, full headwind of reasonable and constant velocity. 3.14.2 Factors affecting safe landing performance Apart from the pilot's physiological condition, airmanship, experience and capability — and the aircraft's weight and balance condition — landing performance is limited by the following constraints, all of which should be carefully assessed — both within the pre-landing procedure and at the flight planning stage — to establish whether a safe landing is viable. Generally most of the engine effects and other constraints affecting take-off performance, covered in section 11.3, have no significant effect on landing performance — except, with both tailwheel and nosewheel aircraft, the inertial effect of the cg position. However, when a landing attempt is aborted, then any of those constraints may be present during the initial go-around. Demonstrated landing distance. Landing distance is the total distance required to clear an imaginary screen, 50 feet (or 15 metres) high, placed before the airstrip threshold; then touch down and bring the aircraft to a halt with normal braking — in nil wind conditions. It should be borne in mind that the manufacturer's 'demonstrated' landing distance has been achieved by a very experienced test pilot in very favourable conditions, during the type certification tests. The landing distance required by the average recreational pilot may be considerably greater. Airfield dimensions and slope. The usable length of runways or strips must be ascertained, as well as the degree of slope — both with and across the direction of landing. Landing downslope will reduce deceleration and lengthen the ground roll. Slope across the landing path makes the touchdown and subsequent ground roll more difficult to control. At a 'one-way' airstrip a combination of airfield slope and rising terrain at the high end necessitates landing upslope, no matter what the wind direction. Airfield surface and surrounds. A short, dry grass or rough gravel surface might decrease the ground roll by 10% compared to that for a smooth, sealed surface. Wet or long grass might decrease the ground roll by 30%. However, there is a possibility that a wet surface can induce aquaplaning/hydroplaning, which adversely affects braking and/or can result in a ground loop (where the aircraft suddenly swings through 180° or more with probable undercarriage and propeller damage). Frosty grass provides little friction, so be wary in early morning shadowed terrain. Long grass and weeds can catch a wingtip, resulting in a ground loop. A soft or waterlogged surface might greatly decrease the ground roll but will increase the possibility of the aircraft tipping over during the ground roll, or may delay — or even prevent a take-off — if such is attempted during the landing ground roll. The location and height of constructed obstructions, trees and local topography must be assessed. Airfield density altitude. This is a critical factor that is often not correctly assessed. High density altitude has a major effect on the approach speed (i.e. the true airspeed is significantly greater than the indicated airspeed), and thus the ground speed at which the aircraft touches down and the length of the subsequent ground roll. High density altitude also affects the aircraft's climb-out performance if the landing is aborted. Re-read the section on high density altitude. Wind velocity and turbulence. Wind strength, direction, downflow, gust intensity, surface turbulence and the potential for wind shear events are normally the major considerations in landing performance. Read the micrometeorology turbulence module, but particularly the section on 'lee wind downflow and eddies'. You should also read the CASA Advisory Circular 'Safety during take-off and landing'. This is an abridged web version for recreational aviation. The pilot-in-command of an aircraft must assess all the foregoing factors and conditions to ascertain the total distance required for obstacle clearance and landing, judge if the landing can be conducted safely and ascertain a safe go-around route if the landing should need to be aborted. All the foregoing assumes that the height of the cloud base allows sufficient visibility, and appropriate terrain and obstruction clearance within the circuit. The problem for the less cautious pilot — if the airfield conditions are found to be unsuitable — is that an eventual landing is mandatory and, if flight planning is poor, there may be no acceptable alternate airfield within range. 3.14.3 The standard circuit pattern For at least the past 65 years, a standard procedure has been adopted for any piston-engine light aircraft approaching to land at a non-controlled public airfield. This procedure is called the standard circuit pattern and is adopted by convention rather than laid down by regulation. Following the pattern requires that an aircraft should track over at least three legs of a rectangular course aligned with the runway or landing strip that is most into-wind. Turns, once established within the circuit, will all be in the same direction, usually to the left unless terrain or ground habitation dictate otherwise. The downwind leg will be flown at moderate speed (adjusted to avoid overtaking preceding aircraft) and at a constant height — normally 1000 feet above the airfield level is recommended, but some primarily ultralight airfields may have a lower standard circuit height. And, of course, the aircraft must be operating in visual meteorological conditions [VMC] — clear of cloud and in sight of the ground at all times, if at or below 1000 feet agl. Check the visual meteorological conditions for aircraft operating under the visual flight rules. Consistency The height of the circuit is particularly important for ultralight pilots. Ultralight engines and associated systems are not renowned for their reliability and the circuit height should be sufficient that, following power loss, an aircraft flying a reasonably tight circuit has every chance of gliding to a safe landing area on the airfield. Pilots should adopt their own personal circuit procedures, to be used wherever possible; the principle being that consistency improves performance. Do not automatically apply the procedure utilised at a training airfield when operating elsewhere. The skills involved can only be assimilated by repeated practice at many airfields — not by reading books or web pages. Consistency is the key. Every circuit and landing should be performed to the best of the pilot's ability; such consistency makes the occasional difficult landing easy. The diagram below (adapted from the Sydney Basin Visual Pilot Guide, courtesy of the Australian Civil Aviation Safety Authority's Aviation Safety Promotion program) demonstrates the full routine for a piston-engine aircraft inbound for landing at a public airfield. The routine 1. The first stage is an overflight at a height not less than 1500 feet agl (preferably with Local QNH set, but if this is not obtainable, use Area QNH) to determine the airfield serviceability, the surface wind direction, the runway/strip being used by other traffic and confirmation of the circuit direction; or if no other traffic, to select the strip to be used. While in the circuit, keep monitoring the relative position and the movements of other traffic at all times. Note that the 'circuit area' is taken to cover the area within a radius of three nautical miles from the 'airfield reference point'. Assume that the latter is the runway intersection. If the airfield is unfamiliar, the overflight also provides the opportunity to examine the circuit area for safe escape routes from each runway following a late go-around. Also check the area for suitable forced landing sites and associated hazards should the engine fail during a go-around or after take-off. See the Coping with Emergencies Guide. 2. The second stage is to manoeuvre so that a let-down from 1500 feet is commenced on the 'dead' side of the active runway, tracking close and parallel to that runway. This is the upwind or into-wind leg. The first and second stages provide the opportunity to carefully check the airfield area and boundaries for hazards — animals, power lines and other wires, ditches, obstructions, and to ascertain the whereabouts of other traffic in, or joining, the circuit and to be seen by them*. All manoeuvring should be done so that the airfield activities always remain in sight; i.e. don't turn away for a short time and then follow with a reversed turn onto downwind. *The official term for this latter procedure is 'unalerted see and avoid', but it has its limitations. See the Australian Transport Safety Bureau research report 'Limitations of the see-and-avoid principle'. The report was first issued in 1991 when mid-air collisions in Australian general aviation averaged about one per year but collisions have increased slightly since then. Most — or nearly all — general and powered recreational aviation mid-airs occur in the circuit area, generally when one aircraft descends into another from behind. 3. When circuit height is reached and the upwind end of the runway has been passed, choose an appropriate position to turn onto the crosswind leg so that there will be no conflict with traffic on the crosswind and downwind legs, and to achieve optimum traffic spacing. You are now entering the traffic side of the circuit. Watch for aircraft joining the circuit on crosswind and for aircraft taking off; ensure that you provide adequate clearance. Maintain circuit height and, allowing for drift, track at 90° to the runway. 4. Turn 90° onto the downwind leg at an appropriate distance past the runway (after checking for aircraft joining the circuit on the downwind leg), check the crosswind drift against selected landmarks and adjust heading to track parallel to the runway, perform the appropriate downwind cockpit checks, and hold altitude and appropriate traffic spacing. Set power and trim the aircraft to maintain an airspeed that allows time to plan the landing without unnecessarily delaying other traffic — probably around 1.7 × Vso. Note: although we call these legs 'upwind', 'crosswind' and 'downwind', they are only nominally named so, because the surface wind is unlikely to be closely aligned with the 'into-wind' runway — particularly with a single strip — and the wind at circuit height might vary considerably from that at the surface. 5. Planning time! Pick an intended touchdown target on the airstrip. This should be far enough into the strip so that an undershoot on approach will still allow normal roundout and touchdown on the runway, or an overshoot on approach will still allow ample runway to bring the aircraft to a halt. For all ultralights and most light aircraft, the latter requirement is probably inconsequential for most runways at public aerodromes. A touchdown target maybe 400 feet from the threshold is about the norm; never target the beginning of the runway or strip for touchdown. Now choose another point, say 200 feet back from the touchdown target towards the threshold; this is the aiming point. Of course, it may be difficult to identify such positions at a featureless airstrip; also, the figures will vary according to the aircraft's drag characteristics in the landing configuration. We are presuming here that we are operating at the average recreational aviation airfield where the strip length may be 2000–3000 feet. It can be a little embarrassing for the light aircraft pilot who touches down 400 feet past the threshold of a 6000 feet runway and then has to taxi a kilometre to the next exit. At a certified aerodrome, the runway centre-lines are 100 feet [30 m] long with a 100 feet gap in between, and the 'piano keys' which normally mark the threshold are also 100 feet long. There should also be touchdown marks at 500 feet [150 m], 1000 feet [300 m] and 1500 feet [450 m]. 6. At an appropriate distance past the aiming point, turn 90° onto the base leg, and hold airspeed but reduce power so that a descent is started during the turn. Lower the first stage of flap if so equipped. Reduce airspeed (but not less than 1.5 × Vso), and trim. The time spent flying base leg is most important, as it provides the opportunity to: set up the aircraft in the approach attitude; establish a power and flap setting (and trim) for the required rate of descent; check for conflicting traffic both airborne and on the ground and particularly any traffic on a straight-in approach or very wide circuit; assess the crosswind component along the landing path; decide the touchdown technique appropriate for the conditions; and review the pre-landing checks. Hold an accurate heading on base to carefully monitor drift, comparing the wind velocity at that height with the surface wind indicated by the windsock(s). A significant difference between the two indicates wind shear will be encountered during the final approach — this may erode the safety margin between the approach speed and Vso, or cause other difficulties. Never be tempted to fly a semi-circular base with a short final approach — it is very poor airmanship and negates all the safety check features of the square base leg. It may be that preceding traffic conditions preclude a turn onto base at the optimum position — in which case you must reduce speed and/or extend the downwind leg further downwind; maintain altitude; and delay the start of descent, and some actions, until the aircraft is well into the base leg or even established on final approach. 7. Start a 90° descending turn onto the final approach so that, on completion of the turn, the aircraft is lined up with the extended notional centre(line) of the landing strip. During the turn, be aware of the reversal height phenomena and confine external scanning to the intended flight path and to the check for conflicting aerial traffic particularly ahead of and behind you. Watch for aircraft on or near the runway; if in doubt about safety initiate a go-around. If satisfied with the initial approach, then lower full flap (if the wind speed is fairly high, then partial flap may suffice), adjust airspeed to the recommended final approach speed [Vref] and re-trim. Once stabilised in the final approach, control the airspeed and the rate of descent with small movements of flight controls and throttle. The power setting should be such that it allows small power reductions, or power increases, in order to maintain the approach path. This can't be done if the approach is set up with the engine at idle power. In addition, the thrust response is not that effective from an idle setting and, for many aircraft, an approach at idle power will entail a high sink rate, which may be difficult to manage. Also, an idle power approach tends to over-cool the engine and may promote carburettor icing, both of which may result in high power not being available when needed — such as in a go-around. If flying an aircraft with a low approach speed into a relatively high wind************* 8. Continue tracking down 'final', whilst correcting for the crosswind component, and watching the position and apparent movement* of the aiming point relative to the windscreen. Then at 50 feet or so, substantially reduce the rate of descent, reduce thrust to zero, touchdown and roll-out until it is safe to turn off the landing strip. If so equipped, and in a nosewheel aircraft, brakes may be applied to slow the aircraft during the latter part of the roll-out — but only if the aircraft is moving in a straight line on a firm surface and the elevators are raised to keep excess weight off the nosewheel. In a tailwheel aircraft, be very wary of any brake application during the roll-out. The braking systems in ultralight aircraft are generally only provided for light use in ground manoeuvring. * If the aiming point appears to be moving up the windscreen you are undershooting (too low) and will touch down before the target. If the aiming point appears to be moving down the screen you are overshooting (too high) and will touchdown past the target. If it appears to be motionless in the screen the approach slope is good and touchdown will be close to the target. The foregoing presumes that all of the runway is visible through the windscreen during the final approach. However, there are some aircraft where the forward visibility over the nose is inadequate at approach speeds and special techniques, such as side-slipping, may be required. Variations on joining the circuit The previous discussion outlined the full circuit pattern that should be adopted when inbound to an unfamiliar airfield. However, when inbound to a familiar airfield of which you are aware of the current runway in use and its serviceability, it may not be necessary to overfly the airfield, and the circuit may be joined anywhere on the green path; i.e. on the upwind, crosswind or downwind leg. Downwind joins are normally made at a 45° angle from outside the pattern. You should not join the standard circuit on base or final — the red shaded path in the diagram. When joining crosswind or downwind, you should already be at the circuit height. Note that only the pattern of the standard circuit is fixed. Its dimensions; e.g. the length of the downwind leg or its distance from the runway, are variable. It is good practice to fly a nice, tight circuit. This also allows a forced landing to be accomplished safely on the airfield if power is lost. However, for operational reasons, not all aircraft will fly a standard pattern or even base their circuit on the same runway. The turning radius of regular passenger transport [RPT] aircraft is too large to conduct the normal circuit pattern, so they perform either a 'circling approach' or a 'straight-in approach'; the latter being much safer for RPT aircraft. Agricultural aircraft reloading at a public airfield tend to use a runway and circuit pattern which best suits the job conditions. CASA have produced two new (2010) advisory publications to support procedures and provide guidance on a code of conduct to allow greater flexibility for pilots when flying at, or in the vicinity of, 'non-towered' aerodromes; i.e. airfields in Class G airspace. These Civil Aviation Advisory Publications are: CAAP 166-1 'Operations in the vicinity of non-towered (non-controlled) aerodromes' and CAAP 166-2 'Pilots responsibility in collision avoidance in the vicinity of non-towered (non-controlled) aerodromes by 'see and avoid'. Please read the combined CAAP 166-1/166-2 document. Note that the 'ultralight' term used in the CAAPs when recommending a 500 feet circuit height, refers only to those RA-Aus aircraft which have a normal cruising speed below 55 knots, or thereabouts. CASA have also produced an online interactive learning tool titled 'Operations at, or in the vicinity of, non-towered (non-controlled) aerodromes' which is now available at casaelearning.com.au/M02/index.htm. 3.14.4 Non-standard circuits Special procedures for joining on final apply at non-towered aerodromes. Aircraft joining for a straight-in approach should be established on the straight-in approach heading by not less than three nautical miles from the airfield; in addition, the aircraft's landing lights and anti-collision lights must be switched on. The straight-in approach option is available to any aircraft (though not recommended) but should only be utilised by aircraft whose approach speed is much higher than the norm; e.g. RPT aircraft. An aircraft on a straight-in approach must give way to aircraft already reported established on base or final approach. The straight-in approach is often made on the longest runway, not necessarily the into-wind runway. Joining on the base leg is also available but not recommended. Refer to the procedures section of the VHF radiocommunications guide for the standard broadcasts on the CTAF. Operational need and the pattern flown The following extract from an older Australian Civil Aviation Safety Authority Advisory Circular AC 91-220(0) concludes that "Safety rules permitting, the pilots of each type of aircraft will want to fly the circuit pattern most suited to the aircraft and the type of operation. Pilots have to give and take relevant information and exercise tolerance and consideration if varied circuit flight paths and experience levels are to be accommodated safely." Extract from that draft AC 91-220(0) regarding operations at non-controlled aerodromes. The principal factors or elements relating to operations in VMC are: The type of operation — agricultural, pilot training, air transport Type of aircraft Wind speed and direction Number of runways Obstructions and topography in the vicinity of the aerodrome Built-up areas and local noise sensitivity Number of aircraft Other activities — parachuting, glider flying, flight training Whether all aircraft are radio-equipped and proximity of controlled airspace and low-level operations Non-communicating traffic and non-compliant traffic. There can be varied operational needs and manoeuvres conducted at a non-controlled aerodrome: Skilled pilots will often want to make smaller circuits than pilots under training or with low recency Larger air transport aircraft are expensive to run, and minutes saved make straight-in approaches an attractive proposition Helicopters are not restricted to normal circuit patterns and generally operate to stay clear of fixed-wing circuit patterns Pilots doing actual or practice instrument approaches will often make straight-in or abbreviated approaches to a landing or to a missed approach point on an instrument runway, or will elect to join the circuit from overhead a navigation aid via the most convenient turn to the runway in use Agricultural pilots conducting local deliveries may prefer to do a contra or a low-level circuit, or make straight-in approaches on a cross runway (expect any legitimate manoeuvre that will speed up delivery rates) Parachuting and glider tug aircraft may make steep descents into the circuit area Ultralight pilots generally prefer to make low, small circuits, and to overfly terrain with potential for a safe forced landing Gliders require winching or towing, often use parallel runways and/or contra circuits, and are committed to land from the time they enter the circuit Trainee pilots require relatively large circuits, don't have reserve capacity to cope with unusual manoeuvres by other aircraft , and can easily be forced to abandon their preferred flight path by other aircraft, including those on normal manoeuvres. Though a little out=of-date the complete CASA draft advisory circular 91-220 (0) makes useful reading and has been provided on this site as 'Operations at non-controlled airfields. 3.14.5 Final approach slope and duration Large aircraft on the final approach to the runway normally descend along a documented path which is inclined at about 3° to the horizontal and aligned with the runway. All Instrument Landing Systems [ILS] are based on this 3° (or 5%) approach slope; the term glideslope is usually accepted to refer to the approach slope in such systems. Most of the secondary aerodromes in Australia are equipped with the ground aid Visual Approach Slope Indicator systems [VASI], or something similar; these day and night optical indicator systems also utilise the 3° glideslope. Thus for larger aircraft, the approach technique is to intercept the glideslope some distance from the runway threshold and to maintain a consistent airspeed and rate of descent throughout the straight-in approach. The rate of descent necessary to maintain the glideslope is controlled by slight power changes and depends on the effect of wind; i.e. the ground speed. The rule of thumb for the required rate of descent in feet per minute along a 3° slope is the ground speed in knots multiplied by 5. This is just another application of the 1-in-60 rule. One knot = 100 feet per minute, so if the ground speed is 120 knots (12 000 ft/min) the rate of descent required to maintain the slope is 12 000 × 3/60 = 600 ft/min. If a 20 knot wind reduces the ground speed to 100 knots, the rate of descent required reduces to 500 ft/min. Maintenance of the glideslope and direction (the track over the ground should follow the extended runway line) are the critical needs in a precision approach. Thus it is also necessary to assess the crosswind component of the wind velocity and make the necessary heading adjustment to compensate for drift. Light aircraft approach slope and speed For light aircraft approaching at a ground speed of, say 50 knots, the 3° slope is not really practical as the rate of descent required would be only 250 ft/min. This extends the time spent on final which, in turn, tends to back up the traffic in the circuit. Also, maintenance of a documented approach slope is not a critical need in an approach that is not instrument, GPS or ground aid oriented. Glideslope management for light aircraft entails a bit of mental arithmetic to either: calculate the rate of descent required plus monitor the VSI — if fitted, or if more comfortable with a particular rate of descent, calculate the ground distance necessary between aiming point and final approach point (see below). Light aircraft generally use a steeper approach slope — maybe around 6° which, at 50 knots ground speed, would require a rate of descent of 500 ft/min. The rule of thumb for the rate of descent to maintain a 6° slope is the ground speed in knots multiplied by 10 equals the rate of descent in feet per minute. The manufacturer's recommended final approach speed [Vref] chosen for light aircraft in normal approaches is usually not less than 1.3 × Vso, possibly 1.5 × Vso for low speed aircraft. (The slower the aircraft, the greater the effect of atmospheric turbulence.) The planned rate of descent is usually established by pilots as one they are comfortable with, at the final approach speed. Airspeed and the rate of descent, at a particular flap setting, are controlled by small adjustments in attitude and power. Sideslipping adds another dimension to the approach angle. You should review 'forces in a descent' and the 'lift/drag ratio'. For a normal approach it is important to hold — and trim the aircraft into — the recommended approach speed without adding any extra 'safety factor'; the safest approach and landing will be achieved at that recommended airspeed. An allowance for wind gusts should be added if necessary, or 2–3 knots may be added in significant crosswind conditions (see below). The duration of the final approach then depends on the height from which 'finals' are commenced and the planned rate of descent. In a normal approach, the final approach is usually started at about 400–500 feet agl with a chosen rate of descent around 400–500 feet per minute; thus the time on final should be about one minute. The over-the-ground distance covered during final approach depends on the duration, the approach airspeed and wind velocity. Taking a low momentum ultralight approach as an example, if the turn onto final is completed at 500 feet agl, the rate of descent is 500 ft/min, the approach speed is 50 knots and the headwind velocity is 10 knots, then the ground speed is 40 knots (4000 ft/min), the duration is one minute and the final approach must start about 4000 feet from the aiming point. The lower the ground speed (as with a stronger headwind), the lesser the ground distance must be between start of final and the aiming point, otherwise you end up conducting a low 'drag it in' approach. This is not good energy management,as it is both low and slow — and totally reliant on engine power to keep you out of trouble. It is probably unwise to use full flap when confronted with high wind speed on the approach because, under the conditions just described, you will be flying the back of the power curve with significant power required to balance the increased flap drag; it is better to choose a flap setting that provides a higher CL without a substantial increase in CD. Final approach point Having chosen the rate of descent, the height at which the final approach will commence and estimated the wind velocity, then sometime during the downwind leg the pilot must determine the ground position that marks the final approach point — the point where the turn from base onto final will be complete. The position at which the preceding 90° descending turn — from downwind onto base — should be commenced is determined by that final approach point and the wind velocity. Presuming that the wind direction at circuit height is roughly aligned with the landing direction, then the higher the wind speed, the earlier the turn onto base must be started. Allowing for crosswind When the aircraft is flying the upwind, downwind or base legs, the allowance for drift — in order to maintain a tidy rectangular track around the circuit — is always accomplished by assessing the necessary wind correction angle or crab angle and adjusting the aircraft's heading so that the aircraft 'crabs' along the required ground line. The crab method is also used on final approach, particularly in larger aircraft, to adjust for the crosswind component. Rudder, rather than aileron, is used to make small adjustments to the aircraft heading. The crab method is the most comfortable for passengers. However, the forward slip method is probably easier to manage in some light aircraft if the crosswind component becomes significant on final approach. The main thing in handling crosswind is to ensure that the aircraft is not moving sideways at touchdown; i.e. the longitudinal axis is aligned with the direction of forward movement and that direction should preferably be aligned with the runway or strip. Sideways movement at touchdown stresses the undercarriage and may prompt a violent swing. In an ultralight, if the crosswind component is becoming a bit extreme you can always reduce it a bit by landing diagonally (i.e. edge to edge) across a (wide) runway or strip. The crosswind component and its relativity to aircraft speed will vary as the aircraft descends due to the decreasing wind gradient and the reductions in aircraft speed. Particular care should be taken when landing upslope, as the wind speed might drop off very rapidly near the surface, due to the blanking effect of the terrain. 3.14.6 Flare, touchdown and ground roll During the final approach, the aircraft should be descending towards the aiming point. Maybe a few seconds before it will fly into that point, the aircraft is 'flared' so that the aircraft's attitude is smoothly changed — from the nose-down attitude of the approach to a nose-high attitude for landing. During this 'round-out' transition period, power is smoothly reduced to idle, or near idle, and the aircraft's vertical speed is reduced from maybe 400–500 ft/min to practically zero. At the same time, its forward speed is also reduced from the approach speed to about 1.15 × Vso, plus any wind gust allowance. Because the aircraft is turning in the vertical plane, wing loading will increase during the flare so stall speed during that period will be slightly above Vso. If the flare elevator pressure is excessive, the aircraft will 'balloon'; i.e. the nose will point skyward and airspeed will drop off very rapidly in a (very) short climb unless immediate corrective action is taken. At the end of the flare manoeuvre, the aircraft should be flying level just above the surface and decelerating as it approaches the touchdown target. An aircraft close to the surface will be in ground effect and the decreased induced drag will mean that the rate of deceleration slows; i.e. the aircraft will tend to 'float'; the higher the ground speed, the longer the float duration, and the greater the chance of encountering some difficulty due to wind gusts, lulls or shifts. If you approach with a tailwind, the aircraft will seem to float forever. The drag from fully extended flaps will increase deceleration and reduce float. The duration of the float will be minimised by an approach at the correct airspeed plus a firm, smooth round-out and power reduction. The touchdown airspeed chosen by the pilot depends on wind conditions, and there are two touchdown options. The usual technique is for the pilot to ease the main wheels onto the surface while finally closing the throttle, touching down lightly while the aircraft is in a somewhat nose-high attitude but still above Vso — a 'wheeler' landing. This technique is always used in unfavourable wind conditions. Sometimes, rather than the pilot flying the aircraft onto the surface, the aircraft might be held in that attitude just above the surface until airspeed decays and the aircraft lands itself. At touchdown — in a taildragger only — some forward pressure may be applied to the control column until the speed decays below Vso, pegging the aircraft down with the reduced aoa so that it cannot lift off again, while airframe drag and wheel friction are slowing the aircraft. A nosewheel aircraft should never be allowed to touch down nosewheel first, or the nose and main wheels together, as wheelbarrowing may result. The nosewheel should be held off the surface during the roll out until the aircraft slows, and then gently lowered, rather than letting it drop down of its own accord. Keep the aircraft aligned with rudder. The alternative technique is to 'hold-off' the touchdown by gradually increasing control column back pressure, and holding the wheels a few centimetres above the surface as the airspeed decays. Recalling the formula: Lift = CL × ½rV² × S, you can see that in this technique the pilot is preventing the aircraft from touching down,and holding lift constant by increasing CL as V² reduces. When close to the stalling aoa and the airspeed is near Vso, the pilot stops increasing back pressure and the aircraft sinks, alighting smoothly in a nose-high attitude. This technique is particularly suitable for tailwheel aircraft — but only in favourable wind conditions. The object is to touch down simultaneously on the main wheels and tailwheel; i.e. a 'three-point' landing, without the aircraft sinking very far. When using this technique in a nosewheel aircraft you must not allow the nosewheel to thump down when the main wheels touch. If the 'crab and kick' technique is used to compensate for crosswind, then the aircraft's fore and aft axis must be finally aligned with the direction of movement by kicking the rudder just before touchdown occurs; good timing is necessary. After touchdown maintain runway alignment with rudder. A refinement, requiring a very fine touch on the controls, is to crab until very close to the runway then gently lower the into-wind wing so that the main landing gear on that side contacts the runway, then using rudder, pivot on that wheel to align with the runway centre-line. Similarly, if the forward slip method is used, then touchdown is made on the into-wind main wheel before the airspeed decays below Vso. The weight should be kept on that wheel until the aircraft slows at which stage the other wheel will contact the surface. If a nosewheel is interconnected to the rudder pedals for ground steering — and it remains connected even if the weight is off the nosewheel strut — then the nosewheel will be deflected in flight by the use of rudder. Touchdown of a deflected nosewheel must be avoided so the rudder must be in the neutral position before the nosewheel is lowered to the surface. During roll-out in crosswind conditions, the into-wind aileron is raised to prevent that wing from lifting — if gust-effected — and direction is controlled with rudder. In a taildragger, the pilot must be prepared to counter the inertial effect of the centre of gravity position. Unless there is a good reason for doing so — a touch-and-go landing, for example — flaps should not be raised until the aircraft has reduced to taxiing speed or turned off the landing strip. If there is some distance to taxi, then before turning off, it is safe practice to move to the side of the runway from which you will turn, to leave room for another aircraft — just in case. Soft field technique If the airfield is soft then the technique is to minimise the weight on the main wheels at touchdown, gradually transferring the weight from wings to main wheels as the aircraft slows. The approach is normal, using full flap if available, and the aircraft is flared as normal for a reasonably nose-high attitude — but a little power is applied just before touchdown, as you feel for the surface. Be prepared for the aircraft nose to pull down hard as the wheels sink — the same nose-down pitch will happen if touching down in long grass, particularly if it is wet. Remove the power smoothly, do not touch the brakes (a locked wheel will not ride over any obstruction), hold the control column well back and keep the aircraft moving until you attain firm ground. The rebound effect The rebound effect following a heavy, main wheel landing differs between tailwheel and nosewheel aircraft. A taildragger's cg is behind the main wheels, while a nosewheel aircraft's cg is in front of the main wheels. Thus the inertial effect combined with the reaction forces generated by the tyres and shock-absorber gear of a taildragger — acting vertically ahead of the cg — will tend to rotate the aircraft nose-up during any rebound, thereby increasing the aoa and thus lift. The aircraft bounces high, induced drag increases and a series of bounces or even pilot-induced oscillations could be initiated. The possibility of a stall with wing-drop is high. The opposite effect occurs with a tricycle gear aircraft; the rebound effect will tend to rotate the aircraft nose-down, reducing aoa and lift and thus bringing the nosewheel closer to the surface; the initial bounce is mild and any subsequence bounces might be described as skip-bounces; the chances of wheelbarrowing increase. Recovery from a bad bounce is probably best achieved by going around if safe to do so, otherwise by adding a little power and easing the aircraft into the proper condition for a smooth landing. 3.14.7 Going around A go-around is a decision to abort the landing and climb straight ahead (perhaps to rejoin the circuit on the crosswind leg), and involves a transition period between the descent phase and subsequent climb. A go-around decision might be taken at any time during the final approach, the flare, or sometimes even after initial surface contact. If the decision is made late, then the transition period might be a critical time for the pilot because of the low energy status of the aircraft and its low-speed flight characteristics. For lower-powered aircraft, the go-around technique requires a full, smooth application of full-throttle power to arrest the descent (followed by checking carby heat control to cold air and, if fitted with variable pitch propeller, ensuring pitch is set to maximum rpm), then maintain level flight while building kinetic energy or perhaps even trade some height for faster acceleration. Only commence the climb-out when Vx, Vy or an intermediate climb speed is attained. If the aircraft is low when the go-around decision is made and power is applied, then continuing to descend so that the aircraft can be accelerated in ground effect will provide some additional airspeed should that be considered safe and desirable. There may be occasions when a cooled (or iced-up) engine fails to respond to the throttle being opened in a go-around following a throttled-back glide approach or a practice forced landing approach. (The same lack of response may occur if the throttle is opened too rapidly.) Consequently, the pilot should be careful not to raise the nose before, or at the same time as, opening the throttle because — if the engine doesn't respond, there will be no increase in thrust to balance the substantially increased drag; sink rate will increase and the wings will approach the critical aoa. Generally, the aircraft will pitch up with full application of power and it should not be necessary to apply very much control column back pressure, but raise the nose AFTER the engine has responded properly. If the aircraft is equipped with flaps, then the flap retraction procedure for a go-around should be specified in the pilot's operating handbook. Generally, to avoid dangerous sink, flaps should be raised slowly in stages — and only when a positive climb rate is established, and obstacles are cleared — then finally cleaned up when a safe height is reached. Some aircraft will not climb at full throttle with full flap deflection (this particularly depends on gross weight, cg position and density altitude but perhaps is further complicated by rising terrain) in which case it is necessary to reduce to an intermediate flap setting during the transitional stage of the go-around, while applying just sufficient control column back-pressure to negate the sink. If climbing with approach flaps extended, the aircraft's attitude in pitch may differ substantially from the normal climb attitude. If a go-around decision is made when the aircraft is on the ground with full flaps extended, then set take-off flap before applying full throttle. If the aircraft has a retractable undercarriage (and unless the pilot's operating handbook states otherwise), then do not to raise the gear until the climb is well established and other more vital procedures can be completed — without distraction from the primary task of maintaining aircraft attitude and airspeed. There is always the possibility of the aircraft sinking to the surface if it is low when flaps are first raised, or mistakenly stowing all flap instead of raising the undercarriage. Pilots must be able to select and adjust flap positions, trim positions and undercarriage control without looking around the cockpit. At a public airfield, regulations require the aircraft to maintain runway heading until 500 feet agl. However, there may be a local convention that suggests aircraft track a climb-out path that follows safer terrain, in case of engine failure. Density altitude will severely deplete an aircraft's go-around performance. If high density altitude is combined with high gross weight and a short or uphill strip, then a go-around may be impossible. The reasons for a go-around from base or final approach might be: a perceived traffic conflict the landing area fouled an unstabilised approach or one that requires too many major changes in throttle setting an excessive sink rate on final, which may be evidence of downflow turbulence the approach is just too fast, too high or low, way off the landing line, or just confused. Go-arounds at or after touchdown are usually prompted by multiple bounces arising from a high rate of sink at first contact. Any time you have to pour on power to regain control of the aircraft, it is probably mandatory to then go around — provided there is sufficient remaining runway, there is a safe climb-out path ahead and the aircraft is not swinging. Many airstrips used by recreational light aircraft are just that — a strip lined by trees, scrub or soft sand. So, if the aircraft has swung away from the strip alignment, a go-around under those conditions may be unsafe. It may be preferable for you to make an early decision on the type of accident you may have by closing the throttle, establishing a reasonable aircraft attitude, holding tight and preparing for some relatively minor aircraft damage. It is better to hit the obstacles when groundborne rather than airborne; see 'Engine failure after take-off'. Here is an extract from an RA-Aus incident report: The pilot reported that ... "aircraft touched down in slow wheeler landing, bounced in semi-stall condition and yawed through 90 degrees. Full power applied, power lines were 140 metres away and line of 50 foot trees were 180 metres (away). Aircraft climbed at maximum angle of climb but neared the stall as trees got closer. Downwind component and high humidity didn't help the situation. Aircraft cleared the line of trees but then stalled and clipped a tree behind the first row of trees." In the preceding report you might perhaps substitute 'high density altitude' for 'high humidity'. The decision to go around must be executed positively as early as possible — don't be indecisive and don't start a half-hearted go-around attempt. Here is another extract from an RA-Aus incident report: "The pilot was practicing short landings and low power approaches. Just before the point of flare he decided to go around and applied power. After the aircraft had begun to gain altitude he decided to land ahead on the remaining runway. Again unhappy with the situation at the point of flare he decided to go around and reapplied power. At this point the left wing dropped and the aircraft slewed off the centreline and struck a sapling growing off to the side of the runway. The pilot was not injured but the aircraft suffered major damage." Elevator trim stall At each stage of the approach, the aircraft should be properly re-trimmed to maintain the desired airspeed at the current cg position and selected flap configuration. The elevator trim tabs exert quite a large control force at flight speeds. With full flap deflection on the approach, some aircraft may need quite an amount of nose-up trim; under these conditions, the application of full power following a go-around decision may induce a very strong nose-up movement — exacerbated by the elevator trim setting — and this attitude change must be anticipated by the pilot. If the pilot is slow in applying forward stick pressure and adjusting the elevator trim, the pitch-up may result in a highly dangerous 'elevator trim' stall. A similar situation may occur when conducting touch-and-go landings. On the other hand, if a lot of nose-down trim has been applied during the approach to landing, that also may cause difficulties on a subsequent go-around or touch-and-go if the pilot neglects to re-trim to the appropriate take-off setting. Read 'Running out of runway' in the July–August 2002 issue of 'Flight Safety Australia'. A go-around undertaken when the aircraft is low in energy has a much greater risk profile than a normal runway take-off, and thus must be conducted with considerable care. With the engine producing high power and the aircraft's attitude changing, the engine effects — propeller torque, gyroscopic precession and P-factor — will also be evident in a go-around. These effects must be anticipated and compensated for. Any turns conducted at a low energy level must be gentle and coordinated. See the safety brief 'Loss of control in low-level turns' and read this RA-Aus accident investigation report. 3.14.8 Short field techniques Planning for a short field landing is started with the airfield check in the flight planning stage. This is where the pilot ascertains airfield dimensions, slope, surface conditions, obstructions and hazards, plus forecast meteorological conditions. The next step is to calculate the aircraft's take-off distance under those conditions. If the calculations show ample margin for take-off, then landing — for a flap-equipped aircraft — should be okay, as it usually (but not always) requires a shorter distance for landing. It is best to do the complete landing distance calculation, factor in all the known conditions — but assume nil wind — then multiply that calculated distance by 1.5 to allow room for error. If the result is greater than the distance available, then landing at that field is unsafe. If the distance available is greater than 1.5 times but less than perhaps twice the calculated distance, then the approach and landing should be planned using short field landing techniques. The problem with airfields considered short for recreational light aircraft is that they are often poorly engineered, single, private strips and generally surrounded by obstructions. Some are built on rising ground, so that the landing can only be done uphill no matter what the wind velocity. Once committed into the final approach, the feasibility of going around is very doubtful. Do not plan to land at an airfield that is both short and one-way — it is venturing into the realm of gambling, not flying; and is most unlikely to be acceptable to the aircraft insurer. The following is an extract from an RA-Aus incident report: "The aircraft was being landed on a one-way strip with a tail wind. When it became apparent that the aircraft was not going to touch down in time, power was applied in an attempt to go around. The aircraft could not climb enough to clear some obstacles in its path so was turned to avoid them and, after clearing a shed, it struck a tree and came to rest. The pilot, who described himself as 'very, very lucky', was not injured even though the seat belt was torn from its mounting in the impact." It is not just the physical length of an airstrip that must be considered: under high density altitude conditions, many 'normal' airfields become 'short' and those same high density altitude conditions may preclude a go-around. Wire hazards Short airstrips seem to have an affinity for power cables to be strung across the runway ends — though I am aware of one private airstrip where the power supply to the house is strung across the middle of the runway, supported at each side by two poles. Remember, it is the wire the pilot didn't know was there — or knew was there but didn't see — that all too often brings an aircraft to grief. The following is an extract from an RA-Aus incident report: "The pilot departed his airstrip for one owned by a friend about eleven nautical miles away. Approaching from the west and about 1.5 nm from the threshold he began a gentle descent, passed over a set of power lines, then flew over a second set of lines, reducing his speed to 55 knots to set the aircraft up for landing. At this point he noticed the owner of the airstrip standing about halfway along the strip. The pilot, judging his approach to be OK, was contemplating whether he might need a slight application of power to flare on the uphill threshold when the aircraft struck a third set of power lines. The lines caught the propeller, exhaust and undercarriage, causing the aircraft to decelerate and strike the ground in a vertical attitude before coming to rest inverted. The pilot suffered minor bruising to the head and the aircraft was substantially damaged. The pilot involved supplied a list of 'points to ponder': 1. He had driven to the airstrip and inspected it from the ground three weeks previously. 2. He had previously landed on the airstrip from the east. 3. On the day before the accident he had overflown from the west in a different aircraft and then landed from the west. 4. The western end of the airstrip is in a localised low area and the poles carrying the power line were both obscured, one by a house and the other by trees." All of this indicates that pilots should be extremely wary of marginal airstrips and never carry a passenger into such situations. Perhaps many should be avoided, as they allow little margin to cope with micro-meteorological events that cannot be forecast — such as gusty crosswinds or lee downflows — and where a landing is really just a demonstration of pure bravado, perhaps with a dash of stupidity. Technique Getting into a short field requires accurate energy management (i.e. height and speed); firm, smooth controlling; and a properly calibrated ASI. You will have to choose a touchdown point that is closer to the threshold than normal, commence the flare a fraction later than normal and ensure the approach airspeed is slower than normal so that the approach angle is steeper than normal — particularly once clear of obstacles, and float is minimised by the aircraft being placed firmly on the surface soon after round-out. The steeper approach allows for obstacle clearance while still achieving the earlier touchdown, and also it keeps a little more potential energy of height in hand. If at any time you are not happy with the approach, then initiate an early go-around using the correct go-around technique. Be decisive — don't wait to see if you can recover the situation. Also, if you have made two missed approaches, then perhaps it's time to go elsewhere. Here is another extract from an RA-Aus incident report: "The pilot had made two downhill into-wind approaches to a short sloping strip but was unhappy with the speed of the aircraft and decided to approach downwind/uphill. As the aircraft touched down about 50 m along the strip he decided to go around and applied full power. The aircraft cleared a fence at the top end of the strip but then dropped a wing and landed heavily, collapsing the nosewheel and damaging the right main wheel. The pilot was not injured." Choosing a bug-out point and 'escape' route Short field landings require a little more preparation, starting with a slower initial overfly at 1500 feet agl and turn onto the upwind leg. During this period, find something that will clearly mark a point about halfway along the selected landing path. This will be the go-around point; i.e. if the aircraft has not touched down when it reaches this point, following the flare, then a go-around will be decisively initiated. You should be aware that an airstrip that is much smaller than those you are used to may prompt the tendency to scale down the circuit and the illusion that you are too low on final approach. See 'Runway illusions' in the March–April 2000 issue of 'Flight Safety Australia'. You must plan an escape route for the go-around from that bug-out point, and determine whether the aircraft will have the climb performance to clear any obstacles and high terrain on that route. Be aware that terrain slope discerned from 1500 feet agl is likely to be under-estimated. Also take into account that atmospheric conditions near the surface may not be what you expect. If you have any doubts — do not attempt a landing; that little voice telling you 'maybe this is not a good idea!' is understating the situation. You should plan not to touch down at the first pass, but to initiate a go-around before the flare and above obstruction height. This gives an opportunity to explore the final approach without any commitment to land. The low pass also provides a chance for a closer look for obstacles at the runway ends, a check of the surface condition and cross-slope, and to run off any wildlife. Landing routine 1. Follow the normal routine on the downwind leg, except fly it a little slower, lower partial flap and reduce to normal approach speed before commencing the descending turn onto base. This will provide more time to hold heading on base so as to carefully check for wind shear, which may further erode the safety margin between the reduced approach speed and Vso. If shear is indicated, a decision must be made whether to continue or to abandon the landing attempt. On base, reduce airspeed, lower full flap and keep the aiming point in sight. During the descending turn onto final, use a touch more power to balance the increase in induced drag and maintain the lower airspeed. Remember, during the approach, it is essential to re-trim the aircraft at the required airspeed after each flap and power change. 2. As early as possible after being established on final approach, reduce airspeed to the short field approach speed recommended in the Flight Manual. If that doesn't exist, use an airspeed that is at least 1.2 × Vso and a low power setting — you will tend to control airspeed with elevator and descent with power. The lower the power setting, the greater the sink rate. Remember, you will be flying the back of the power curve and the power setting used should be enough that there is ample reserve for a go-around if needed. Watch for apparent movement of the aiming point in the windscreen, and adjust power or airspeed to hold that point motionless. Also watch the top of the highest obstacle along the approach path. If the vertical distance in the windscreen between the top of the obstacle and the aiming point is widening, you should clear the obstacle; if it is narrowing you may not clear it. Start reducing the power when clear of obstacles. A suitably experienced pilot in a non-flap equipped aircraft can steepen the approach by sideslipping, but not with an inexperienced passenger as the manoeuvre can be a little frightening. 3. The slower approach speed means there is need to accurately maintain airspeed within 2 or 3 knots without continuous reference to the ASI, hence the need to accurately adjust trim. (All the foregoing presumes that the ASI accuracy, or variance from the stated Vso, has been calibrated). During the round-out, there will be a need to apply a slightly greater back pressure on the control column. This results in a consequent increase in wing loading and a further reduced margin between the accelerated stall speed and the airspeed, plus a greater tendency to balloon. Also, the possibility of an elevator trim stall following the application of full power, if a go-around is initiated, is more likely. Further reading The online version of CASA's magazine 'Flight Safety Australia' contains articles relating to landing that are recommended reading. Look under 'Take-off and landing' in the 'Further online reading' page. Signals that are essential to know When radio communication cannot be established by airfield control, there are five internationally recognised light signals that may be used to advise air and ground traffic at that airfield. A hand-held signalling lamp is used to direct the signal at an individual aircraft. The signals are a steady or a flashing green; a steady or a flashing red; and a flashing white light; as below: Light signals Directed at aircraft on the ground Directed at aircraft in flight Steady green — authorised to take-off if the pilot is satisfied that no collision risk exists Authorised to land if the pilot is satisfied that no collision risk exists Flashing green — authorised to taxi if the pilot is satisfied that no collision risk exists Return for landing Steady red — stop Give way to other aircraft Continue circling Flashing red — taxi clear of landing area in use Do not land Airfield unsafe Flashing white — return to starting point on airfield Before landing, it is essential to check the ground signal square usually located adjacent to the white primary windsock. The displayed ground signals denote the airfield operational state. Aerodrome is unserviceable, do not land. A cross or crosses displayed on a manoeuvring area denote unfitness for use. Aircraft operations are confined to hard surface runways, aprons and taxiways only. See AC 139-06 January 2011 'Use of restricted operation (dumb-bell) ground signals Gliding operations are in progress (and gliders have priority right of way) Wind direction indicators or windsocks Wind direction, variability and strength is usually assessed by observing the airfield windsocks — these indicate the direction and variability, and may provide some idea of the wind speed a few metres above the surface. Indication of wind speed will vary with the type of windsock. CASR Part 139, a Manual of Standards for Australian licensed aerodromes, requires one standard white windsock as the primary wind direction indicator located near the signal area, plus additional standard windsocks (yellow in colour) placed near runway thresholds. The standard cone dimensions are 3.65 m [12 feet] long, tapering in diameter from 900 mm [36 inches] at the opening to 250 mm [10 inches] at the exit. The standard light fabric sock indicates a speed of 15 knots or greater when it becomes horizontal in dry conditions, and about 7–8 knots when drooping at 45°. A wind speed above 2–3 knots is usually sufficient to provide a direction indication. Some windsocks may be colour-banded (red/white or orange/white) for higher visibility. Not all operators of aircraft landing areas comply with CASR Part 139, so a variety of wind direction indicators exist. They may be made from heavier materials; e.g. canvas, in lengths from 1.5 m to 7 m. Their positioning and condition range from useless to good. Wind speed indications may vary considerably from those of the standard. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.13.1 The take-off sequence The full take-off sequence starts at pre-flight planning and concludes when: the aircraft is established in the climb configuration at an appropriate threshold height at the best rate of climb airspeed or a suitable enroute climb airspeed with the recommended power setting. The pre-flight planning, weather and airfield check, aircraft inspection, fuel quantity and quality check, engine warm-up and check, taxiing checks, pre-take-off checks and radio procedures are all part of the full pre-flight procedure and of good airmanship; and must be conducted for every take-off — even if you just contemplate doing a quick weather check flight. Take-off procedures and techniques vary according to aircraft type: seaplane or landplane, tailwheel configuration — tractor or pusher engine; nosewheel configuration — tractor or pusher; flap equipped; canard configuration; delta-winged; powered parachute; or weight-shift aircraft. Some procedures should be specified in the pilot's operating handbook for that aircraft. In this module, we will look at the common factors to be considered in the execution of the take-off for the normally configured, three-axis, nosewheel or tailwheel aeroplane. There are differing take-off procedures or techniques, or combinations thereof, applicable to particular airfield conditions: normal take-off short field take-off soft field take-off. The take-off sequence is varied according to prevailing conditions, but it usually has at least three parts: the initial ground roll, where the essentially landborne machine is accelerated to a lift-off speed selected according to the airfield conditions. Aerodynamic drag and rolling friction retard acceleration and the distance required to reach lift-off speed is dependent on atmospheric conditions. It is also inversely proportional to the achievable acceleration — i.e. a 20% increase in acceleration (×1.2) will decrease the distance to 83% (1/1.2=0.83) of the original. Conversely the ground roll is proportional to the lift-off speed squared — i.e. increasing the required lift-off speed by 10% (×1.1) will increase the distance 1.21 times. lift-off followed by a short transition period where the aircraft is accelerated by keeping induced drag to a reasonable level, possibly in ground effect (i.e. while held just above the surface), until either a minimum take-off safety speed (Vtoss) or the selected CAS for best rate of climb (Vy), or the best angle of climb (Vx), is reached. the climb-out, tracking the runway heading, to a safe threshold height where the pilot's options are less restricted, possibly 300–1000 feet above ground level [agl], and where airspeed can be increased to an appropriate enroute climb speed. Regulations forbid turns away from the extended runway line until the aircraft is 500 feet agl. However, at many smaller airfields, local custom may prescribe a climb-out path that provides greater safety in an engine failure event. 3.13.2 Factors affecting safe take-off performance Apart from the pilot's condition, experience and capability, take-off performance is limited by the following constraints, all of which should be assessed carefully within pre-take-off procedure to establish whether a safe take-off is viable. Aircraft weight and balance. The critical nature of aircraft weight and balance at take-off has been highlighted in the 'Weight and balance' module, and should be reviewed. Standard take-off distance [TOD]. TOD should always be expressed as the total distance required to accelerate from a standing start, and clear an imaginary screen 50 feet (15 m) high. The ground roll is that first part of the TOD where the aircraft's weight is partly or fully supported by the undercarriage; sometimes people incorrectly refer to the ground roll as the TOD, ignoring the fact that the distance covered from the lift-off point to climb to 50 feet may be longer than the ground roll. It is known for an under-powered aircraft to be able to lift-off but then be unable to climb out of ground effect. TOD is officially expressed as the take-off distance required [TODR] to clear the 50-foot screen. These standards require that the operating conditions associated with a particular TODR will be specified in approved aircraft take-off performance charts. These conditions are pressure altitude, temperature, runway slope and surface, and wind velocity. CAO 101.28, an airworthiness certification requirement for commercially supplied, amateur-built, kit ultralights states in part (at paragraph 3.6): "The take-off distance shall be established [by the manufacturer] and shall be the distance required to reach a screen height of 50 feet from a standing start, … appropriate to a short dry grass surface … [The] aeroplane [should reach] the screen height at a take-off safety speed [author's emphasis] not less than 1.2 Vs1 … Take-off charts … shall schedule distances established in accordance with the provisions of this paragraph, factored by 1.15." CAO 101.55 has much the same wording but specifies 1.3 Vs1 as the take-off safety speed and FAR Part 23 is similar. 'Short dry grass' means grass less than 100 mm long that is not wet. Unless the manufacturer's take-off performance figures are published as an approved performance chart within the aircraft's flight manual or comparable document, then such figures should be treated as unverified sales claims. In the absence of any specified conditions in an unapproved performance chart, assume that sea-level ISA, nil wind and smooth, dry runway are the basis for the published data. If the manufacturer's performance charts only provide data for the aircraft at maximum take-off weight then, for a recreational aircraft, a reduction of 10% in TODR for each 50 kg the aircraft's weight below MTOW is probably a reasonable estimate. Stopping distance required. The distance required to reach flight speed, and then bring the aircraft to a halt, should be known. It may be necessary to abandon the take-off soon after lift-off, due to doubtful engine performance or other event — this is particularly important in short field or 'hot and high' take-offs. If take-off and landing distance (over a 50-foot screen) charts are available then the total distance needed to take off, abandon take-off at 50 feet, land and bring the aircraft to a halt is the sum of the charted take-off and landing distances required. Airframe condition. An airframe in a battered or dirty condition, or which has unnecessary or non-standard accoutrements, will increase drag and retard acceleration, lengthen TODR and reduce climb performance. Engine age, condition and operating temperatures. An engine that is incapable of producing its rated power will reduce acceleration, lengthen TODR and reduce climb performance. The engine manufacturer's instructions regarding warm-up procedures should be followed, to ensure appropriate temperatures and pressures are established before the engine is subject to the stresses of take-off power; otherwise the potential for an engine failure after take-off is greatly increased. Check carefully for any warning signs or sounds during the full power ground roll. Never continue with the take-off if there are any doubts. Propeller condition and pitch. Chipped leading edges or scored blades, apart from being dangerous due to the possibility of delamination or fracture, will adversely affect thrust output. Blade pitch at a coarse setting — a cruise setting — will reduce acceleration and climb performance. Tyre pressure. Under-inflated tyres increase the rolling friction, decrease the acceleration and add perhaps 10% to the ground roll. Airfield dimensions and slope. The usable length of runways or strips must be known, as well as the degree of slope. Taking off upslope will reduce acceleration and lengthen the ground roll because thrust must also overcome a force equal to the aircraft weight × the sine of the angle of slope, in addition to the drag and rolling friction. The ground roll will increase by about 15% for each 2% of upslope. Runway slope can be measured by taking an altimeter reading at each end, dividing the elevation difference by the runway length (in feet) and multiplying by 100 to get the approximate slope percentage. Airfield surface and surrounds. A short. dry grass surface or rough gravel surface might add 10% to the ground roll compared to that for a smooth, sealed surface. Wet or long grass might add 50% to the ground roll. A soft or waterlogged surface might double the ground roll. Surface water and/or wet grass can lead to aquaplaning and loss of directional control; the effect of frost is similar. The height of obstructions and local terrain must be known. Airfield density altitude. The density altitude is a critical factor that is often not correctly assessed, and has a major effect on engine output, propeller performance and lift generated. Thus it affects acceleration, TODR and climb performance to such an extent that on 'hot and high' airstrips an aircraft may be incapable of safe take-off and climb-out. Read section 3.4 'High density altitude'. Wind velocity and turbulence. After weight and balance plus density altitude, the major considerations in take-off performance for a properly maintained aircraft are then wind strength, direction, gradient, downflow, gust intensity, surface turbulence and the potential for wind shear events. Please read 'Surface gusts or low level wind shear' in the 'Wind shear and turbulence' module. The diagram indicates possible cumulative effects of some take-off conditions on TODR. But as explained in section 11.6, the take-off distance required can be much greater. The pilot in command of an aircraft must assess all the foregoing factors and conditions to ascertain the cumulative total distance required for take-off and obstacle clearance, and then judge if the take-off can be conducted safely. The golden rule is "If you have ANY doubts, don't fly". The most favourable conditions for optimum take-off performance at MTOW are: a pilot who follows the rules and the recommended procedures a certificated aircraft in very good condition and fitted with a 'climb' or variable pitch propeller a surface that is dry, smooth and level — or with a slight downslope a low density altitude; i.e. low elevation and low temperature a smooth, full headwind at ground level of reasonable and constant velocity sufficient separation is maintained to avoid aircraft wake turbulence. You should not only be concerned that the take-off is conducted safely, it should also be accurately controlled — beginning with taxiing — so that alignments, headings, attitude and airspeeds — the 'numbers' — are properly maintained throughout. The take-off should take advantage of the aircraft's and engine's maximum rate of climb capability to reach the threshold height — and it should look well executed to an informed observer standing behind the aircraft's take-off point. In addition, you must have pre-established plans to safely cope with partial or total power loss, occurring at any stage of the take-off sequence. See 'Engine failure after take-off' and 'The turn back, possible or impossible — or just unwise?'. There are web versions of two CASA Advisory Circulars on this site: Operations at non-controlled airfields and Safety during take-off and landing. Both these documents should be read in conjunction with this module. 3.13.3 Engine/propeller effects and ground effect There are some engine effects, plus aerodynamic and inertia phenomena, which will be noticeable at take-off. However, both their existence and the extent of their effect are dependent on the configuration of the aircraft. Tailwheel aircraft are particularly subject to these phenomena, which can cause difficulties to any pilot who is inexperienced in the slow-speed handling of such aircraft. Ultralight aircraft also tend to have a much higher power-to-weight ratio than their general aviation counterparts. For example, at MTOW, the two-seat 110 hp Cessna 152 and Piper Tomahawk both weigh 1670 lb and have a power loading of 15 lb/hp; whereas a two-seat amateur built aircraft acceptance category ultralight equipped with an 80 hp engine will have a power loading of 12.5 lb/hp, and only 10 lb/hp if fitted with a 100 hp engine. A single-seat CAO 95.10 ultralight fitted with just a 60 hp engine will have a power loading of 11 lb/hp. The lower the power loading, or the higher the power-to-weight ratio, the greater and faster the reaction will be to the engine/propeller effects. The helical slipstream The propeller blades produce a rotating slipstream tube with a diameter equal to that of the propeller disc and a helical effect that increases as forward speed increases. If the propeller rotates clockwise, when viewed from behind the aircraft, the slipstream tube will also rotate clockwise. Where the engine is mounted in the nose (as with the Jabiru), then the slipstream will rotate clockwise around the fuselage; anything mounted below the fuselage will experience increased pressure on the right side (from the slipstream striking it at an angle) and anything mounted above the fuselage will experience higher pressure on the left side. The significant surfaces mounted above the fuselage are the fin and rudder, and the increased pressure on their left-hand side will tend to push the tail to the right; i.e. in nil wind conditions, the aircraft will want to swerve to the left — particularly in the early stages of the take-off run when the slipstream counts for practically all the airflow around the fin and rudder. The swing direction would be reversed for aircraft where the propeller rotates anti-clockwise. Full application of compensating rudder may be required at the start of the ground roll. The helical effect lessens as the aircraft accelerates (because the angle at which the slipstream meets the vertical surfaces lessens and also the rudder becomes increasingly effective), so rudder pressure should be decreased as the take-off roll progresses. Slipstream effect is not so apparent in the landing ground roll because normally the throttle is closed. However, if the engine is mounted above the fuselage, the rotating slipstream tube will be higher relative to the fin and rudder, and the swing effect may be lessened or reversed; aircraft with a pusher engine mounting are subject to the same effect. Before you fly any aircraft it is advisable to determine which way the aircraft will swing, and how to control the swing. The helical slipstream will also meet the horizontal stabiliser at an angle but the resulting effect is difficult to determine or distinguish. When a tailwheel aircraft has all wheels on the ground, as in the early part of the take-off ground roll, the slipstream may be deflected by the airfield surface so that the effect on the fin and rudder may vary between the tail-down and tail-up positions. Propeller torque effect The reaction torque of a propeller rotating under power attempts to rotate the aircraft about the propeller shaft. Of course, it is prevented by the resistance of the wings and undercarriage. However, at the beginning of the take-off run, the torque may be sufficient to increase the friction on one tyre and thus cause the aircraft to pull towards that side. The effect is there in the early stages of take-off but may not be apparent as such, because it reinforces the swing tendency initiated by the helical slipstream. (The propeller torque on some very high-powered, piston-engined fighter aircraft has been such that at full power the aircraft tended to hop sideways down the runway. In such aircraft, the engine was not opened up to full climb power until airborne, unless it was carrying a very heavy armament load.) Gyroscopic precession effect Any external force, which tends to alter the direction of the angular momentum axis of a spinning gyroscope, causes the direction of the axis to move (precess) 90° to the applied force and in the direction of rotation. A fast-rotating propeller disc acts as a gyroscope spinning in the lateral plane, its moment of inertia (the resistance to a change in angular velocity about the propeller shaft) is proportional to the propeller mass and the disc diameter squared. When the aircraft's attitude in pitch or yaw is changed rapidly the aircraft applies a torque to the propeller disc and the propeller's reaction is an equal and opposite moment or force applied to the aircraft. But the gyroscopic precession effect causes the direction of that moment to move (precess) 90° to the applied force and in the direction of propeller rotation. For example, if the aircraft's attitude is pitched up the upper rim of the propeller disc is forced back while the lower rim is pushed forward. The precession moment is moved 90° clockwise* to the applied force so the upper rim becomes the disc's right side (looking from the rear) and the reaction moment is directed to the rear tending to yaw the aircraft to the right, i.e. during the period the aircraft is being rotated about its lateral axis the gyroscopic precession effect is also trying to rotate the aircraft about its normal axis. Similarly if the aircraft is pitched down the precession effect prompts a yaw to the left. Conversely if the aircraft is strongly yawed to the left the nose tends to pitch up; if yawed to the right the nose tends to pitch down. There is no gyroscopic precession effect when the aircraft is rolled about the longitudinal axis. *Note: assuming a clockwise-rotating (viewed from behind the aircraft) tractor or pusher propeller. The magnitude of the gyroscopic moments induced by the rotating propeller are dependent on the rate of change in aircraft pitch or yaw, the rotational speed of the propeller and its moment of inertia. The precessive forces are transferred via the shaft to the propeller speed reduction unit or direct to the engine crankshaft, bearings, crankcase and mountings. Sport and recreational aircraft generally have a high power-to-weight ratio and the engines apply unusually high rpm to the propeller. The most prevalent example of the gyroscopic effect in such aircraft is in the early stages of a taildragger's take-off run should the pilot shove the control column forward to raise the tail and accelerate. At this stage airspeed is low so the aerodynamic forces generated by the airframe are also low and have a decreased ability to counter the gyroscopic effects. The pitch down causes the aircraft to yaw to the left so the pilot must anticipate this action by applying compensating rudder as the tail is lifted. Even ground manoeuvring may induce unfavourable gyroscopic effects – swinging the aircraft around with a burst of power plus rudder/brake places high loads on the propeller shaft. For an example of the possible longer term effects on the propeller shaft see 'The Fox story – gyroscopic loads' also see matching engine and propeller. You can read a little about gyroscopic effect in Spitfires and Seafires; the gyroscopic effect is also utilised in some advanced aerobatic manoeuvres in aircraft with powerful engines and large propellers, the Lomcevak end-over-end tumble and inverted spin was the first. P-factor P-factor, or asymmetric disc effect or asymmetric blade effect, occurs when the thrust line is not aligned with the flight path; i.e. when flying with a high angle of attack. As the propeller disc is then inclined to the relative airflow, a down-going propeller blade has a greater component of forward velocity than an up-going blade; thus, the down-going blade generates slightly more thrust than the up-going blade. For a clockwise rotation, more thrust is then generated on the right-hand side of the disc, which again reinforces the slipstream, torque and gyroscopic-induced tendencies for such aircraft to swing left during take-off. P-factor is dependent on thrust and is proportional to forward speed, so it is not a significant factor in the initial part of the ground roll for a tailwheel aircraft, even though the axis of the airscrew disc is inclined to the horizontal; it will become increasingly apparent as the ground roll progresses, if the aircraft's tail-down attitude is maintained. P-factor may also become apparent as higher velocities are reached — just before and after lift-off — if a high aoa is employed at those stages. P-factor may cause the aircraft to yaw when flying level using high power at high angles of attack. P-factor has little or no effect on a tailwheel aircraft during the landing ground roll because, normally, when the throttle is closed no thrust is produced — there is only propeller drag. However, should the throttle be opened suddenly during the ground roll while the tailwheel is on the ground, there may be a prompt P-factor reaction. Inertial effect of centre of gravity position relative to the longitudinal axis If the aircraft's cg is behind the main wheels, as it must be in a tailwheel undercarriage aircraft, then any ground swerve — initiated by the helical slipstream, gyroscopic effect, torque, crosswind, wind gust, deflating tyre or rough ground — will be reinforced by the inertia of the aircraft, applied through the cg position, and tend to pivot around the main wheels. When the cg of the loaded aircraft is in front of the main wheels — i.e. a tricycle undercarriage — the aircraft's inertia will lead to self-correction of the swing, as long as there is no excessive weight on the nosewheel. The cg inertial effect is usually much more likely to cause real difficulties when a tailwheel aircraft is slowing (i.e. on landing) rather than when accelerating. There are circumstances where the cg inertial effect also applies to nosewheel aircraft; see 'wheelbarrowing'. It is very important in such aircraft to identify any departure from the planned heading at a very early stage of the 'swing' and take prompt, corrective action — but not to the extent of over-correcting. The pilot must recognise the swing, stop it, correct the heading and then halt the correction. Over-correction is exacerbated by a hard, smooth runway surface. A groundloop is a swing that has been accentuated by the inertial effect into a very rapid 180° movement, which often causes wingtip and undercarriage damage, and occurs at speeds between 5 and 25 knots. At low speeds and/or in light winds, the inertial effect is stronger than any weathercocking action. There are occasions when it is necessary for a pilot to induce a groundloop, usually when aborting a take-off and nearing the boundary fence or something solid at speed — or after a misjudged approach and landing. The groundloop is induced by applying full rudder and brake on the appropriate side. The swing effect is exacerbated if a tailwheel aircraft is 'short-coupled'; i.e. the moment arm between the tailwheel and the main wheels (or the fin and the cg) is short, and thus the tailwheel friction moment is less than it might be. Such aircraft swing very rapidly. The inertial effect requires that taxiing techniques for tailwheel aircraft differ from those for nosewheel aircraft. A turn, initiated by rudder or brake in a nosewheel aircraft, will stop as soon as the pilot removes rudder or brake pressure, because the inertial effect is always trying to straighten up the ground path (wind conditions permitting). However, with a tailwheel aircraft, once a turn is initiated the inertial effect will keep the turn going — and possibly tightening — until the pilot takes definite action by using opposite rudder or brake to halt the turn. The inertial effect of the cg position relative to the main wheels is relevant when landing; see the rebound effect. Ground effect In the 'spanwise pressure gradient' section of the 'Aerofoils and wings' module we saw that induced drag was a consequence of lift generation, and the associated wingtip vortices increase the momentum imparted to the downwash. As the centre of each vortex is a little inboard of the wingtip, the vortices also have the effect of reducing the effective wing span, the effective wing area and probably the effective aspect ratio. When an aircraft is flying very close to the airfield surface during take-off and landing, the formation of the vortices is partly impeded by the proximity of the ground, so induced drag is less than normal and the centre of each vortex moves outboard a little with the potential for a little more lift. The phenomenon is ground effect and mainly — because of the drag reduction — produces faster acceleration on take-off (which can be very useful) and slower deceleration on landing (which generally is not useful). It can only occur when the lower surfaces of the wings are much less than one full wingspan distance from the surface. The closer the airborne aircraft is to the surface, the greater the reduction in induced drag. A light aircraft that maintains height with the wing under-surface about one-quarter wing span above the ground, might experience a 30–40% reduction; at low speeds, this would amount to a 15–20% reduction in total drag. A 50% reduction in induced drag might be achieved if the wing height is equivalent to one-tenth of wing span, which may be possible in a low-wing aircraft and if the pilot has a very steady hand. Induced drag is normally a much greater force than the wheel/tyre rolling friction on a smooth, dry surface. If flying in ground effect and utilising maximum available power, then when a disturbance causes the aircraft to lift further away from the ground, the induced drag will be restored immediately with a consequent decrease in airspeed, decrease in lift and substantial sink towards the ground. Similarly, if maintaining a constant low velocity in ground effect (i.e. not accelerating, which is poor energy management practice but can readily occur in an underpowered or overweight aircraft, or when attempting take-off in high density altitude conditions) the aircraft may not break out of the ground effect because as the control column is pulled back, the induced drag increases, velocity slows, lift decreases and the aircraft sinks back into ground effect. If the aircraft cannot be accelerated it may end up tripping over the boundary fence, unless the throttle is closed and the aircraft landed. The effective angle of attack of the horizontal stabiliser is also affected, mainly by the changing angle of the wing downflow. This might be evident as an uncommanded but slight pitch-up or down when leaving or entering ground effect. The same effect generally applies to seaplanes and amphibians for water take-offs and landings, so 'ground effect' should be more properly termed 'surface effect'. 3.13.4 Calculating density altitude The calculation of density altitude is fully explained in sections 3.3 and 3.4. However, we will run through an example for an airstrip — 'Olly's Folly' — located at an elevation of 2000 feet. Under ISA conditions, the standard temperature and pressure at that height is 11 °C and 942 hPa respectively. We will do density altitude calculations for a cold winter morning in a high pressure system, and a hot summer afternoon in a low pressure trough. Remember that each 1 °C variation from ISA is roughly equivalent to 120 feet variation in density altitude. (a) Cold winter morning: temperature is 0 °C and by setting 1013.2 on the altimeter pressure setting scale we read off the pressure altitude as 1600 feet. (We remember, of course, to then reset the scale to local or area QNH). The temperature of 0 °C is 11 °C less than ISA, so the density altitude variation due to temperature variation is: –11 × 120 = –1320 feet. So, density altitude = pressure altitude ± temperature variation = 1600 –1320 = 280 feet Thus, the aircraft should perform well at take-off — close to its rated sea-level capability. (b) Hot summer afternoon: temperature is 35 °C and by setting 1013.2 on the altimeter pressure setting scale we read off the pressure altitude as 2400 feet. The temperature of 35 °C is 24 °C greater than ISA so the density altitude variation due to temperature variation is +24 × 120 = +2880 feet. So, density altitude = 2400 + 2880 = 5280 feet Thus, the aircraft will perform poorly at take-off — probably at less than 70% of its rated sea-level capability. The following is an extract from an RA-Aus incident report: "I was attempting to take-off in a paddock approximately 140 metres in length. Due to the hot (35 °C) conditions the aircraft did not get enough lift which resulted in the main wheels catching the top wire of the boundary fence. The aircraft was slowed and struck the ground in a nose-down position. The wire snapped allowing the aircraft to bounce approximately 20 feet in the air. I cut the power and landed the aircraft to the left to miss another fence. This caused the left wingtip to strike the ground before coming to a stop. I walked away from the accident." The aircraft manufacturer provided the following information: "... the take-off distance to safely clear a 15 metre obstacle is 213 metres in ISA sea level conditions." Rule of Thumb #1 In the absence of manufacturer-supplied data the effect of density altitude on TODR (for a dry, smooth and level surface) can be estimated: "In nil wind conditions, for each 1000 feet that the pressure altitude exceeds sea level add 10% to TODR, then for each 10 °C that the airfield temperature exceeds 0 °C add a further 10%." e.g. in the 'Olly's Folly' hot day situation, the aircraft manufacturer's standard sea level TODR is 250 m. Pressure altitude is 2400 feet: 250 × 1.24 = 310 m. Temperature is 35 °C: 310 × 1.35 = 419 m TOD. Then add a further 10% margin for random events = 460 m estimated TODR. This is for a dry, smooth and level surface; if the surface is long grass with a 2% upslope then you might have to add another 50% to TODR, making it nearly three times the manufacturer's standard distance! Remember that all the factors mentioned above relating to surface, slope, pressure, temperature, airframe and engine condition are cumulative, and the runway length is finite. Rule of Thumb #2 In the absence of manufacturer-supplied data, the effect of density altitude on maximum rate of climb can be estimated: Let's say our aircraft's manufacturer states the initial Vy rate of climb at sea level in standard ISA conditions is 1000 feet per minute. However, manufacturers' standard sea level rates of climb are usually based on an aircraft in factory new condition, flown by a very accurate pilot in the most benign atmospheric conditions. The manufacturer's standard should be downgraded by a factor that represents an adjustment for general engine, propeller, airframe and other conditions — say 15% — thus the practical rate of climb at sea level in standard ISA conditions should be regarded as 850 feet per minute at Vy. "The practical rate of climb at Vy should be reduced by 10% for each 1000 feet of density altitude." e.g. At a density altitude of 5000 feet, there is a 50% reduction in the maximum rate of climb to 425 fpm. 3.13.5 Effect of wind Wind direction, strength and variability are usually assessed by observing the airfield windsocks — these indicate the direction and variability, and provide some idea of the surface speed. Indication of wind speed will vary with the type of windsock. The Bureau of Meteorology area forecast should provide an indication of the overlying gradient wind. Take-off into wind! There are several reasons why an aircraft, operating from reasonably flat terrain, should normally take-off directly into wind — or as close to that as possible when operating from defined runways or strips. If an into-wind take-off coincides with an upslope runway, then a little calculation should be done to ascertain whether a downslope tailwind take-off is preferable. You may find some 'one-way' airstrips where a combination of airfield slope and rising terrain at the high end mandates take-off in one direction only, no matter what the wind direction. If you intend operating into such strips, check the aircraft insurance policy carefully, because cover may be voided. The ground (rolling) speed for take-off is lower. The airspeed during the ground roll equals the ground speed plus/minus the headwind/tailwind component. Thus, if the aircraft is rolling at 30 knots into a 10 knot headwind, the airspeed = 30+10 = 40 knots. If rolling at 30 knots with a 10 knot following wind, the airspeed = 30 –10 = 20 knots. It is easier to keep straight because of the aircraft's increased directional stability, due to the higher airspeed. The take-off ground roll is shorter. The into-wind climb-out will be steeper and provide better obstacle clearance. (But the rate of climb — i.e. time to height — is not dependent on wind direction.) The vertical wind profile is such that the wind velocity changes encountered during the climb are likely to be an increase in headwind speed, thus providing a momentary increase in lift should any vertical shear be encountered. If the engine should fail after take-off, the aircraft can readily land into wind thus reducing impact force, because the ground speed is reduced quite significantly at light aircraft speeds. However, there are other factors involved; see 'Practice good energy management in the take-off!'. It is safer to conform to an accepted traffic pattern, which is always based on take-off into wind, or as near as runway direction allows. Estimating the crosswind component of the wind velocity When operating from defined airstrips or runways, the chances of the wind direction corresponding exactly with the strip alignment are low; thus, most take-offs have an element of crosswind. Also, local gusts and eddies usually alter the wind strength and direction during take-off. Taking off with a significant crosswind component makes it more difficult to keep aligned with the selected path — because the aircraft will try to weathercock into the crosswind — and increases the possibility of one wing lifting during the ground roll. Lateral forces may stress the undercarriage. All aircraft should have a demonstrated velocity limit for the 90° crosswind component in both take-off and landing. For a very light aircraft, the demonstrated crosswind component limit may be 10–12 knots, beyond which there is insufficient rudder authority to counter any adverse movement. If the crosswind limit is not known, you can assume that it is less than 25% of Vso. (FAR Part 23.233 requires that all aircraft have safe handling characteristics with a direct crosswind component not less than 0.2 Vso.) There are also various techniques to be learned for positioning the ailerons, elevators and rudder — depending on aircraft configuration, wind strength and wind direction — while taxiing and during the ground roll. While taxiing, the aircraft will always tend to weathercock into wind and there are techniques for taking advantage of that when turning in breezy conditions. Be aware that, due to the high cg and narrow wheel track, all light aircraft are fairly unstable when turning while taxiing. Turns made at speeds much above walking pace may result in a wingtip ground strike. Easy calculation to determine the crosswind component Having determined take-off direction and estimated the wind velocity: 1. Estimate the wind angle; i.e. if you intend taking off towards the north and the wind is coming from the north-east or north-west, then the wind angle is about 45°. 2. The crosswind component is the windspeed multiplied by the sine of the wind angle. However, a reasonable approximation of the crosswind component is made if you multiply the wind angle by 1.5 and apply the result as a percentage (to maximum 100%) of the wind speed. e.g. Wind speed 15 knots, wind angle 45°: Crosswind component = 45 × 1.5 = 67.5% of 15 = 10 knots If the angle was 30° the crosswind component would be about 7 knots. 3. If the wind angle is 60° or more, consider the full wind speed as the crosswind component; i.e. wind speed 15 knots, wind angle 60°, then crosswind component = 15 knots. Estimating the headwind or tailwind component In some crosswind take-offs, you may need to estimate the headwind or tailwind component of the wind velocity. The headwind or tailwind component of a crosswind is not the wind velocity minus the crosswind component — the square of the headwind or tailwind component equals the square of wind velocity minus the square of the crosswind component. Easy calculation to determine the headwind or tailwind component Having determined take-off direction and estimated the wind velocity: 1. Estimate the wind angle; i.e. if you intend taking off towards the north and the wind is coming from the north-east or north-west, then the wind angle is 45°. 2. The headwind component is the windspeed multiplied by the cosine of the wind angle. However, a reasonable approximation of the crosswind component is made if you deduct the wind angle from 115 and apply the result as a percentage (to maximum 100%) of the wind speed. e.g. Wind speed 15 knots, wind angle 45°: Headwind component = 115 –45 = 70% of 15 = 10 knots 3. If the wind angle is 30° or less, consider the full wind speed as the headwind component; i.e. wind speed 15 knots, wind angle 25°, then headwind component = 15 knots. If the wind angle exceeds 90° from your intended take-off direction then, of course, there is a tailwind component. In which case, use the acute angle that the wind subtends with your take-off direction; e.g. if the wind is from the south-east or south-west when taking off towards the north the acute angle is 45° and the same calculation as above is made to determine the tailwind component. Easy calculation to determine the headwind or tailwind effect on ground roll distance If you know the nil wind take-off ground roll for a particular aircraft, you can estimate the take-off ground roll for various headwind components, with the same airfield surface conditions. The take-off ground roll = the nil wind ground roll × ([lift-off speed –wind speed] /lift-off speed)² For example, if an aircraft has a ground roll of 100 m before reaching the normal lift-off speed of 40 knots, what would be the take-off ground roll into a headwind of 5 knots? The take-off ground roll = 100 × ([40 –5] / 40)² = 100 × 0.875² = 100 × 0.765 = 76 m. What would it be with a tailwind of 5 knots? The take-off ground roll = 100 × ([40 + 5] / 40) ² = 100 × 1.125 ² = 100 × 1.265 = 126 m. As you can see, there is a significant difference (50 m) in ground roll even in light winds. If the wind speed components involved were 10 knots, the ground roll would be 56 m into a headwind and 156 m with a tailwind. 3.13.6 Take-off procedure In a normal take-off the aim is to arrive at the 50 feet screen height, as quickly as possible, while maintaining optimum flight safety margins (including traffic separation) and without undue stress on the undercarriage. Normal take-off — nosewheel three-axis aircraft Let's imagine a nosewheel undercarriage aircraft (having lined up in the chosen direction and ensured that the nosewheel has trailed in the fore and aft position) just starting its take-off run, with the throttle being smoothly advanced to maximum power. The airframe will be in a level attitude and, if the wings have a 4° angle of incidence, the angle of attack will also be 4°. The aircraft's total weight is supported on the main wheels and the nosewheel. The rudder will be held deflected in a position to counter the initial slipstream and torque effects, with applied rudder pressure reducing as acceleration progresses. Ground roll. As the ground roll accelerates — because thrust is greater than the rolling friction plus total drag — the airflow velocity increases. At a speed perhaps 20% below Vs1 the elevators should have enough authority so that a little back pressure on the control column will provide sufficient up-elevator to raise the nosewheel from the surface, and increase the aoa by 2 or 3 degrees to 6 or 7°. This may slow the acceleration rate slightly but the reasons for getting the nosewheel off the ground earlier than really necessary — and holding it there — are: the nosewheel strut is the weakest part of the undercarriage and more susceptible to damage from a rough surface the support of the aircraft weight is then shared between the main wheels and the wings rolling friction, being proportional to weight on the wheels, is reducing as lift is increasing the ride is smoother on the main wheels only raising the nose a little reduces the possibility of stone or weed damage to the propeller without excessive deterioration of the view forward. Also, if on a smooth runway and you try to hold the nosewheel on the ground, by increasing forward pressure on the stick as the speed builds, you run the risk of wheelbarrowing. This is where the wings are generating sufficient lift (particularly if take-off flap is set or you are conducting a 'touch and go' landing and take-off) so that the weight on the main wheels is reduced (or they even lift-off slightly) and an abnormal part of the aircraft's weight is riding, and pivoting, on the nosewheel. Under these conditions, the moment arm between the nosewheel and the rudder is very long and the moment applied by the rudder, which is the most effective control at these speeds, is then much greater than normal. Any application of rudder will make the aircraft pivot about the nosewheel rather than the main wheels. The aircraft's cg is now behind the pivot point and the cg inertial effect will make the aircraft behave like a taildragger, but the possibility of a groundloop is greater and the consequences more drastic. On a slippery surface, the aircraft may slide sideways. Wheelbarrowing is a definite no-no on take-off and on landing. As rolling speed builds, so does airspeed and lift. If you allow the aoa to increase beyond 6–7°, a flight velocity may be prematurely reached and the aircraft will lift itself off at an airspeed slightly above stall speed. In this condition, any slight turbulence or mishandling will cause a loss in lift and the aircraft will settle back again, or maybe just one wing drops and it hops about on one wheel. Obviously not a tidy departure; you, not the aircraft, must be in command of the take-off — and you must maintain alignment with some selected reference point throughout the take-off. Rotation. Unless the aircraft manual, or flight school procedures for students specify otherwise, the usual technique is to hold the aircraft at 6–7° aoa until airspeed builds up to a lift-off speed [Vlof] 15–20% above Vs1, then apply further back pressure to rotate the airframe around the main wheels to an aoa of around 12°, and the aircraft will lift off smoothly and commence to climb away with sufficient airspeed in hand to deal with minor turbulence. Anticipate that P-factor effect will cause the aircraft to turn. Do not wait so long that the aircraft flies itself off; you, not the aircraft, should be in command. The increase in induced drag, which is greater than the removal of the rolling friction, will slow the acceleration rate. So, as the initial climb progresses, ease the stick forward until Vy is reached (at an aoa around 8°) and maintain maximum rate of climb at that speed until the planned threshold height is reached. For some aircraft it may be advisable to use Vtoss rather than Vy until a safe height is reached. In gusty wind conditions, it may be prudent to delay rotation until airspeed is perhaps 10% higher than normal. Some nosewheel aircraft may have a tendency to pitch up rather rapidly during rotation and the pilot must be ready to arrest this with forward pressure on the control column. Climb-out. Do not hold the aircraft down to build up speed beyond Vy and then pull up steeply — it displays poor airmanship and is extremely dangerous. Airspeed in a 'zoom' climb like that will drop off very quickly — possibly faster than the pilot can pitch the nose down — which may lead to an irrecoverable departure stall. Take-off procedure may vary a little if the aircraft is fitted with flaps that can be set to a position that provides increased lift without a significant increase in drag. There are other factors involved — see 'Practice good energy management in the take-off!'. Unless stated otherwise in the Pilot's Operating Handbook or engine notes, maintain full throttle until a safe height is reached. The initial climb speed maintained would normally be Vy but if a fixed-pitch cruise propeller is fitted then an airspeed higher than Vy may be more effective. After the initial climb a higher 'enroute climb' airspeed may be the optimum choice to reduce sector time and to maintain engine temperatures within the manufacturer's specified bounds; full power and low airspeed will 'cook' an aero-engine. Also, a lower pitch angle during climb improves forward visibility. Estimating the pitch angle It may be of interest to figure the pitch angle (the angle that the fuselage reference line subtends with the horizontal) during the climb-out. If the aoa is 8° and the angle of incidence is 4° then the fuselage reference line will be inclined at an angle of 4° above the aircraft's flight path. If the aircraft's practical rate of climb at sea level in standard ISA conditions is 850 feet per minute and Vy = 65 knots (or 6500 feet per minute) then the angle of climb (the flight path) is inclined about 8° to the horizontal, so adding the fuselage reference line inclination of 4°, the pitch angle in the climb will be 12°. One event, guaranteed to spoil your day, is the pilot's seat sliding back when the aircraft is rotated and accelerating after lift-off. If your aircraft is fitted with adjustable seats that slide on rails make doubly sure that your seat is locked in a comfortable position before take-off. Also ensure the passenger's seat is locked; she/he may grab at the controls if they find themselves sliding back — that will certainly ruin your day! Incidentally, when initially settling in to the cockpit, make sure that you can comfortably (i.e. without straightening your leg) apply FULL left and right rudder. If you cannot adjust the seat or rudder pedals to achieve this, do not fly that aircraft, because you will not have the full rudder authority provided by the designer. Also there is a danger that, should the aircraft come to a sudden halt with your knee joint locked while applying full rudder, impact forces may damage the knee and hip joint; so, you must be able to apply full rudder with the knee still bent. Normal take-off — tailwheel three-axis aircraft Tailwheel aircraft are subject to all the effects mentioned in section 11.3, and have a fairly predictable mode of behaviour at ground speeds between 5 and 25 knots. They will want to swing and gyrate, and these movements must be anticipated and promptly corrected by the pilot. Ground roll. At the start of the ground roll (again having lined up in the chosen direction and ensured that the tailwheel has trailed in the fore and aft position) the fuselage of a 'taildragger' is naturally pitched up at an angle of maybe 10–12°. Combined with the angle of incidence this means that the aoa at the start of the roll will be close to the stall aoa of about 15°; some aircraft with a high angle of incidence may be past the stall aoa. As the throttle is being smoothly and fully opened at the start of the ground roll, slipstream and torque effects will be at their greatest. Consequently, normal procedure is to start the ground roll with compensating rudder applied, and with the elevators held in the fully up position to put pressure on the tail wheel. The friction of the tailwheel will assist in taming the initial convolutions, particularly if the tailwheel is steerable. However, the high aoa implies consequent high drag and slow acceleration. The tailwheel is the weakest part of the undercarriage, so there is a need to relieve the loads on it as early as possible, particularly if the airfield surface is rough. Also the sooner the generation of lift begins to reduce tyre friction the faster the aircraft accelerates. Thus the requirement is to get the tailplane up reasonably soon so that, firstly, aoa is reduced to 6 or 7° or less; and thus the aircraft is able to pick up her skirts and run. Secondly, the lower the ground speed at which the aircraft's tail is raised, the gentler will be the swing from the ensuing gyroscopic effect. Thirdly, the sooner a near-level (i.e. slightly tail-down) attitude is achieved, the sooner the building P-factor effect is negated. However, remember that gyroscopic effect is also dependent on the rate of change of attitude in pitch, so ease the stick forward rather than shoving it forward. (Propeller surface clearance must be maintained so be careful on non-prepared strips.) Then, as lift-off speed is reached, rotation and climb-out proceeds as for a nosewheel aircraft. Taildragger enthusiasts sometimes refer to the appearance of the aircraft during take-off — when the pilot holds it in a level, minimum drag, maximum acceleration attitude — as being 'on the step'; the term is borrowed from seaplane pilots. Waterborne take-off Although the potential for tyres aquaplaning/hydroplaning — and thus affecting the landing roll — might be considered when landing on a wet runway, surface friction is rarely considered in runway take-offs; however, for seaplanes, water resistance [hydrodynamic drag] dictates the waterborne take-off routine. At rest the seaplane's centre of buoyancy is usually under the forward limit of the aircraft's centre of gravity, while the location of the vortex-creating transverse step in the hull or float/s is usually just to the rear of the centre of buoyancy. After water-taxying to the line-up position, the first part of a seaplane take-off involves getting the aircraft past the 'hump' speed where the aircraft is displacing the maximum amount of water, thereby creating the maximum hydrodynamic drag. (As with air resistance, water resistance also increases in proportion with the aircraft velocity squared.) The throttle is opened, while holding the control column right back, so that the thrust power (including the vertical component of the thrust line) combined with increasing lift from the high aoa quickly lifts the forward portion of the hull/floats above the surface and the hull or floats are 'ploughing' the water, i.e. pushing the water aside. The hump speed might be around 20–30 knots. The second part of the take-off is to to minimise hydrodynamic drag by getting the aircraft operating as a planing hull where it is supported by the hydrodynamic reaction of the water, rather than just pushing the water aside. After attaining hump speed the control column is eased forward to reduce aoa and induced drag and then some back-pressure is applied. So, rather than ploughing through the water and unable to accelerate due to the very high hydrodynamic drag, the aircraft is riding (hydroplaning/aquaplaning) on the deepest part of the hull forward of the step in the hull or floats, total drag (hydrodynamic plus aerodynamic) is greatly reduced allowing the aircraft to accelerate 'on the step' to a speed where wing lift can both break the adhesive action of the water and support total aircraft weight. When on the step the vortices induced by the step break up boundary layer flow and reduce water adhesion to the hull. The aircraft accelerates on the step until rotation speed is reached but unlike a runway take-off, hydrodynamic drag will increase during rotation because more of the hull/float surface is 'wetted' and the aircraft is pushing more water aside. Sufficient thrust must be available to overcome the increase in both the hydrodynamic drag and the induced drag at rotation, otherwise the aircraft can't lift off the water. The 'step-taxying' term describes fast taxying with less than lift-off power. 'Ground' effect is still pertinent in waterborne operations. Short-field take-off In a short-field take-off the aim is to accelerate as fast as possible, be airborne well before the boundary, clear obstacles near the boundary while climbing at the maximum angle of climb, and to maintain reasonable safety margins. Thus we are not so concerned with protecting the undercarriage. The procedure is to maintain a more or less level minimum drag attitude — i.e. 4–5° aoa (with a nosewheel held just above the bumps) throughout the ground roll until Vx is reached — rotate directly to a 12° aoa and climb away at Vx until obstacles are cleared, then reduce aoa to continue the climb at Vy or a higher speed. The ground roll is longer but the acceleration is greater, because rolling friction is normally less than induced drag at a low aoa. You reach Vx in a shorter distance and the TODR is less. The aircraft is subject to all the engine effects but an abnormal P-factor turning tendency should be anticipated after the lift-off rotation. As in normal take-off, the procedure may vary a little if the aircraft is fitted with flaps that can be set to a position that provides increased lift without a significant increase in drag. The recommended flap setting for a short-field take-off may vary from that for other take-off conditions, because the flap position that facilitates minimum ground roll may decrease climb performance. There are some suggestions that flaps should not be lowered to the take-off position until the aircraft is nearing lift-off speed (so the initial acceleration is faster), but the slight advantage provided by this can be dramatically offset by inadvertently lowering the flaps past the take-off position. It is better to set the flaps when doing the pre-take-off checks, when there is time to double-check the selected position. There may be a suggestion that an aircraft equipped with brakes is run up to full power at the start of take-off while holding on the brakes, but generally it is better to smoothly run up to full power while the aircraft is rolling. There is less chance of stone damage to the propeller, and it is easier to prevent a swing developing. Swings and swing correction reduce the acceleration, and it is better to allow time at the beginning of the ground roll to get the aircraft firmly under control. Obviously a take-off into wind is highly desirable, unless runway slope and rising terrain dictate otherwise, and the ground roll should be started as close to the boundary fence as reasonably possible. The procedure described above is for a hard, dry surface or for short dry grass. If the surface is soft or the grass is long and wet, then the rolling friction may exceed the induced drag at medium aoa or the slippery surface may make directional control difficult. In such cases it may be better to get the wheels off early and fly in ground effect until Vx is attained, as in the soft field technique. If there are any doubts about the take-off conditions, then stay on the ground. I suggest you read the article 'Tree's a crowd' in Flight Safety Australia September-October 2002 issue. Soft field take-off Soft field procedures may be applicable to muddy, waterlogged or long/wet grass surfaces. The prime aim in a soft field take-off is to reduce the extremely long ground roll, and become airborne with less than adequate initial airspeed safety margin while utilising ground effect for fast acceleration. The following procedure should not be used in turbulent or gusty conditions, as the possibility of a stall after lift-off is increased. In very soft conditions the usual technique is always to keep rolling; i.e. do not taxi to the take-off position and then stop to do the take-off checks — they should be completed beforehand. When lined up, open the throttle fully and smoothly while holding the control column back. Using a maximum lift flap setting is usually highly recommended. As the elevators become effective, the nose of a nosewheel aircraft will rise. With a taildragger, the elevator pressure should be relaxed sufficiently so that the tailwheel is held off the surface but the aircraft remains firmly in a tail-down attitude. As ground speed builds, start relaxing the back pressure and the aircraft will lift itself (or more likely lurch and stagger) from the surface at its minimum unstick speed [Vmu] and at an aoa very close to the stalling aoa — so, it is vulnerable to turbulence and mishandling. Also, P-factor and slipstream effect may come into play at this time, so it is important to keep the wings level with aileron and stop any turn with opposite rudder to negate any cross-controlled skid. The pilot must then smoothly reduce aoa to 5–6° and hold the aircraft just above the surface in ground effect, so that it accelerates at the maximum possible rate. Gyroscopic effect may be significant during the pitch down to the smaller aoa, which must be anticipated with rudder. The aircraft is rotated after Vy (or Vx if there are obstructions) is attained to break it out of ground effect, held for a few moments to ensure it will accelerate, and then climb-out is commenced. At the initial rotation. the aircraft will slow as induced drag increases substantially and rapidly; firstly because of the restoration of the normal induced drag as it pulls out of ground effect, and secondly because of the increased aoa. The aircraft is likely to sink back to the surface if rotation occurs before sufficient speed is built. The TODR for a soft field take-off will be considerably longer than that for a normal take-off. It is most unwise to attempt take-off from an airfield that is both short and soft. The following is an extract from an RA-Aus incident report: The pilot intended to conduct a trial instructional flight from a grass strip in excess of 250 metres in length. Due to recent rain the strip was soft and several solo take-offs had been carried out, each clearing the fence at the end of the strip by 25–30 m. After some test runs with the passenger on board the pilot's 'gut feeling' was to abandon the exercise but he elected to take-off using a short field technique. The aircraft accelerated until the nosewheel lifted off the ground and then slowed, with the nosewheel sinking back onto the ground. Because he still had what he believed to be sufficient speed in hand the pilot tried to make it over the fence — and didn't. The damage to occupants was minor but the aircraft was a write-off. The pilot identified the cause of the accident as lack of experience in operations from wet fields. In his words the aircraft was 'basically stuck to the field' Coping with significant crosswind During the initial stages of the ground roll in any type of take-off with a significant crosswind component, the aircraft will tend to weathercock into wind and pivot around the main wheels. There are lateral stresses on all wheels in contact with the ground during the roll. The lateral control of the aircraft is then very much dependent on adequate tyre contact with the surface, so if the surface is slippery a crosswind take-off may not be advisable. As the aircraft accelerates, the relative wind velocity (combining the ambient wind velocity, the aircraft's own forward speed and the slipstream velocity) over the tailplane surfaces will have an increasing headwind component and a (relatively) decreasing crosswind component. Thus, it is normal to start the ground roll with a large rudder deflection to counter weathercocking, and decrease the deflection as speed builds. It is usually advisable to also raise the into-wind aileron to prevent the into-wind wing from rising, particularly if gust-induced; the inclined lift vector, because of the rising wing, will tend to turn the aircraft away from the wind. Be aware that if the into-wind aileron is raised while you are countering the weathercocking with rudder, then you must be operating cross-controlled, which will cause the aircraft to sideslip into wind if you should get airborne in that condition. The aileron deflection is decreased as speed builds, but in strong crosswinds it may be advisable to lower the into-wind wing so that the aircraft is rolling just on the into-wind main wheel. The lift vector is then inclined from the vertical and has a lateral component that counteracts the effect of the crosswind; the aircraft line of roll is kept straight by the friction of that into-wind wheel. If the angle is correctly judged, there should be no stress on the wheel. As the aircraft is being lifted off, return the ailerons to neutral and level the wings. To provide an additional safety margin, hold the aircraft on the ground for a higher-than-normal lift-off speed. If conditions are gusty, add 50% of the wind gust speed in excess of the mean wind speed; e.g. if wind speed is 10 knots gusting to 20 knots, add 5 knots to the lift-off airspeed. If the aircraft does become prematurely airborne for any reason then, rather than let the wheels bump down again, hold the aircraft off the ground, accelerate in ground effect and use the soft field take-off technique. After becoming airborne, the aircraft will drift away from the heading, so to mark a tidy and controlled departure, gently turn the aircraft onto a new heading to compensate for the drift and the 'track made good' will follow the extended line of the ground roll — at least until the aircraft reaches 500 feet agl, at which height regulations allow a turn in the circuit direction. It can be that the crosswind either amplifies or reduces the slipstream and other effects. It may be wise to consider taking off in a direction that takes advantage of that counter-effect even if it means taking off with a tailwind component. Also, there is no rule that says you must always take-off aligned with the centre of the runway or strip; if crosswind conditions warrant it, plan your ground roll at an angle across the strip — edge to edge. Traffic separation and wake turbulence Do not commence the take-off roll should until any preceding aircraft using the same runway has crossed the upwind end or commenced a turn, or if the runway is longer than 1800 m the aircraft is airborne and at least 1800 m ahead. However, if both aircraft weigh less than 2000 kg, it is okay to start rolling when the preceding aircraft is airborne and at least 600 m ahead. The runway may be entered following an aircraft landing but the roll should not be started until that aircraft has turned off. The turbulence from the wingtip vortices of aircraft at high angles of attack is particularly strong and a function of aircraft weight. For aircraft taking off, a high aoa is initiated at the start of rotation and continues through climb-out. The wake vortices sink and drift with the wind, and may take several minutes to dissipate. Thus, light aircraft must practise caution when departing behind another aircraft of similar weight, more so if it is a significantly heavier aircraft, as the turbulence will readily roll the aircraft on to its back or worse. When following a very large aircraft note the runway position where the aircraft rotated, wait perhaps two minutes for the wake to dissipate a little, aim to be airborne well before the noted runway position and, where there is any crosswind component, maintain a line along the upwind side of the runway. For a little more information see Aircraft wake vortices in the 'Microscale meteorology and atmospheric hazards' module. Causes of take-off accidents One or more of the following factors commonly cause take-off accidents: exceeding weight and balance limitations failure to set elevator trim at the correct position for the airframe configuration over-controlling during the ground run and at lift-off premature lift-off climbing too steeply after lift-off failure to calculate the TODR to clear all obstacles/terrain and particularly neglecting the effects of high density altitude failure to observe power lines failure to abandon take-off early enough when it is apparent that airfield surface conditions preclude a safe departure using an excessive bank angle in a climbing turn running into the wake vortices from a heavier, previously departing aircraft. Engine failure after take-off [EFATO] Pilots should always be prepared for the possibility that the engine will lose partial or total power during the take-off and climb out; or, for that matter, at any other time during flight. When such an event occurs, the cardinal rule is to fly the aircraft, which initially implies quickly getting the nose down into the right attitude for an appropriate airspeed, either Vbg or Vmp depending on circumstances. Some say the second and third edicts should also be 'fly the aircraft' and 'fly the aircraft'. ( However, if a partial power loss is accompanied by extreme vibration or massive shaking of the aircraft then it is just as important to get the engine completely shut down.) For further information see 'Forced Landing Procedures' in the 'Coping with Emergencies Guide' and 'Engine failure after take-off' in the 'Decreasing your exposure to risk' series. 3.13.7 Precautions when taking off towards rising terrain Take-offs should always be planned so that they do not cause nuisance to others. But it is also prudent to avoid taking off in a direction that takes you close to structures, trees, masts and powerlines unless you are sure that the aircraft will clear them by whatever safety margin you consider acceptable within the existing atmospheric conditions. A take-off towards rising terrain is not something that should be undertaken without a thorough check of all conditions, even if such a take-off has previously been undertaken at a particular location without incident. Density altitude, wind and other conditions may be such that another take-off will result in a 'controlled flight into terrain' incident. Ascertaining terrain height The height of the terrain above the airfield — or more particularly the angle of climb needed to safely clear it — has to be ascertained by whatever means available, to confirm that the aircraft's rate of climb will more than outmatch both the increasing terrain height and the effect of air downflow from the slope. A simple way to judge the angle of climb needed is to extend your arm fully with the fingers bent so that your extended line of sight, including the bottom edge of your little finger, is horizontal. The width of each finger is around 2° and the width of the palm is around 10°. For an example, we will have another look at the 'Olly's Folly' airstrip on that hot summer afternoon. Here there is one grass strip, 1000 feet in length and oriented north-south. Northward, and starting near the end of the strip, the terrain has a 1 in 10 slope rising towards an extensive crescent ridge with an elevation 1000 feet above the airstrip. Using the 1-in-60 rule we can calculate that a 1 in 10 slope equates with an angle of slope of 6°. Ascertaining angle of climb needed We established our aircraft's practical rate of climb at sea level in standard ISA conditions as 850 feet per minute and Vy = 65 knots or 6500 feet per minute. (One knot is near enough to 100 ft/min so to convert knots into feet per minute just multiply by 100.) Then using the 1-in-60 rule we can estimate our aircraft's sea level angle of climb in nil wind conditions, thus: 850/6500 × 60 = 8°. Also note that the ratio of vertical speed to forward speed is about 1:8. But we are not operating in sea level ISA conditions and Vy is only an indicated airspeed, not a true airspeed. TAS is close to 1.5% greater than IAS for each 1000 feet of density altitude, so at our density altitude of 5280 feet TAS is (1.5 × 5.28) % = 8% greater = 65 × 1.08 = 70 knots or 7000 ft/min. Also, our practical rate of climb will be reduced by 10% per 1000 feet density altitude (= 52.8%) to 400 ft/min and the ratio of vertical speed to forward speed has been reduced to 1:18. Using the 1-in-60 rule the angle of climb in nil wind conditions is then: 400/7000 × 60 = 3.4°. Comparing the climb slope with the terrain slope of 6° we can see that it is impossible to outclimb the terrain; in fact the impact point will not be very far from the end of the strip. But what would be the climb angle if we chose to climb at Vx, which should provide a ratio of vertical speed to forward speed 10–15% better than Vy. If Vx, then provided a ratio of 1:15.5, the climb angle using 1-in-60 would be nearly 4°, which would extend the impact point a little further up the slope. Effect of wind on angle of climb A reasonably steady horizontal headwind makes some difference to the angle of climb. Let's say that headwind is 15 knots, which would have the effect of reducing the aircraft's Vy ground speed by 1500 ft/min to 5600 ft/min, so the angle of climb would be 400/5600 × 60 = 4.3°. However, winds that cross over slopes are not horizontal; they may have a substantial vertical component. So the gain, because of the reduction in forward ground speed, may be more than offset by a reduction in vertical speed. In fact, the downflow rate of sink can easily exceed the aircraft's rate of climb, in which case a 'controlled flight into terrain' is inevitable. Re-read lee side downflow in the meteorology section. 3.13.8 Limiting climbing turns during take-off In section 2.8 we discussed the accelerated stall, finding that the airspeed at which an aircraft will stall depends on the wing loading and, as a consequence of providing the centripetal force for the turn, wing loading increases as angle of bank increases. The table in section 2.8 shows that wing loading increases slowly up to a bank angle of 30° — where it is 15% greater than normal — after which it increases rather rapidly — where it is 41% greater at a 45° bank angle. We then concluded that turns involving bank angles exceeding 20–25° should not be made at low levels — including take-off and landing. The wing loading increase in the turn is provided by an increase in CL, which is brought about by an increase in aoa. We also know that the lift coefficient increases in direct relationship to increase in angle of attack. Now what will happen if we are climbing at Vx and decide to quickly turn away from rising terrain or an approaching aircraft, using a 45° bank angle, while still climbing? We know from the table that to maintain a 45° level or climbing turn, wing loading and thus aoa, must increase by 41% and that the aoa at Vx is probably around 12°, so that a 41% increase will take the aoa to 17° and the aircraft will stall. Full power stalls in a balanced climbing turn tend to result in the outer wing stalling first, because of the higher aoa of the outer wing. There will be a fairly fast wing and nose drop (particularly so if the propeller torque effect is such that it reinforces the roll away from the original direction of turn and the aircraft is a high wing configuration) and is likely to result in a stall/spin situation — which any pilot lacking spin recovery experience may find difficult to deal with. If the climbing turn is being made with excessive bottom rudder then the lower wing might stall first, with the consequent roll into the turn flicking the aircraft over. Recovery from a stall in a climbing turn is much the same as any other stall — ease the control column forward to about the neutral position, stop any yaw, level the wings and keep the power on. Even a 30° banked climbing turn at Vx will produce an aoa of 14°, very close to the stall aoa, and provide no margin for even minor turbulence or slight mishandling. The margin you should always have in hand to cope with such events is 3 or 4°. This indicates that when climbing at Vx, turns should not be contemplated. Even when climbing at the Vy aoa (around 8°), until a safe height has been gained, turns should be limited to about 20° to allow an additional margin should wind shear be encountered in the climb-out — and the nose lowered a little for the turn. Further reading The online version of CASA's magazine Flight Safety Australia contains articles relating to take-off that are recommended reading. Look under 'Take-off and landing' in the 'Further online reading' page. Things that are handy to know • The description taildragger is used as a generic term applied to all tailwheel aircraft. However, this is not strictly correct; a true taildragger is an aircraft equipped with a tail skid rather than a tailwheel. Aeroplanes so equipped are usually not fitted with main wheel brakes and they are designed for operation from grass airfields where all take-offs and landings can be made into wind. Such aircraft have little resistance to swinging if operated from a sealed, smooth surface. • Many tailwheel aircraft will have a steerable tailwheel, which improves the aircraft performance during crosswind operations and makes ground handling easier in windy conditions, particularly if not equipped with differential braking. The steerable tailwheel is usually linked to rudder movement and the rudder pedals in some way. But the tailwheel aircraft may have a disconnect feature that allows the tailwheel to fully castor, thereby improving manoeuvring when parking. The steering mechanism may automatically disconnect when weight is off the tailwheel, in which case a spring or other device returns the tail wheel to a low-drag fore and aft alignment. A nosewheel aircraft may be similarly equipped with a steerable nosewheel. Stuff you don't need to know • One of the most successful fighter aircraft of the 1914–18 war was the Sopwith Camel, fitted with a 130 hp rotary engine. The liquid-cooled engines of the day were very heavy and the rotary was designed to utilise air cooling of the cylinders, thus producing a lighter engine for fighter aircraft. The engine rotated around the crankshaft, which was fixed to the airframe. The propeller was bolted to the crankcase so that the engine and propeller rotated as a unit and, because of the flywheel effect, ran very smoothly at normal cruise settings around 1200–1300 rpm. However, as can be imagined, the torque and gyroscopic effects were extreme and such that the aircraft turned very sluggishly to the left but was lightning fast in a turn to the right. If a 90° turn to the left was required it was faster to initiate a 270° turn to the right. In a left turn the aircraft wanted to climb, while in a right turn the gyroscopic effect pushed the nose down. The aircraft was very unstable, hence very manoeuvrable. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.12.1 Hang glider and 'trike' wings and carriages Hang gliding started in 1891 with Otto Lilienthal's first flights; see the history of Australian powered recreational aviation. In 1948 Francis Rogallo, an American aeronautical engineer, experimented in delta-shape flexible wings, which culminated in a project to evaluate his Rogallo parawing concept for suitability as a recovery vehicle for the Gemini spacecraft. That project was finally dropped in favour of parachute recovery, but the technology acquired helped kick-start the modern hang-glider industry. The flexible, swept-wing design provides high lift, reasonable L/D, smaller pitching moments and subdued stall characteristics. The wing is aerodynamically balanced in pitch, because in flight a download is applied at the rear of the flexible wing by a slightly reflexed aerofoil and/or the outer wing sections are washed-out. Longitudinal stability is derived from the reversed centre of pressure (cp) movement — as angle of attack (aoa) increases from the cruising aoa the cp moves backward, which pitches the nose down. The swept-back leading edge provides good lateral stability, although the directional and lateral stability of such wings is also dependent on aoa, being most stable at low speeds. The weight-shift control aeroplane The hang glider technology has developed to encompass heavier and faster, primarily weight-shift control, two-place, powered aircraft with metal-framed, double fabric surface wings; generally called microlights or trikes. Smaller versions may be known as 'nanolights'. See the Airborne Australia website. The powered aircraft carriage (or lightweight cart) — consisting of the pilot/passenger seating, instrument binnacle, pusher engine and propeller mounting and a steerable tricycle undercarriage (from which came the term 'trike') — is primarily suspended, via a streamline-section metal mast, from pitch-and-roll hang point hardware attached to the tubular metal keel of the framed wing structure. The fore-and-aft position of the hang point hardware is ground-adjustable to allow an increase/decrease in the aircraft's trim speed. The pilot's control frame and bar is a fixed part of the wing structure; if the wing is strutted, the inboard end of each strut will be terminated at the control bar. The control bar's neutral position is the aircraft's trimmed level flight position at cruising speed so the aircraft could be flown 'hands-off' the control bar. This arrangement provides direct pilot contact with the wing and the feel for how it is flying. There are no ailerons, flaps or spoilers. Carriages may be an open frame metal structure or a partly or fully enclosed composite pod. Seating for pilot and passenger is usually a very close tandem arrangement. The concept of the carriages and the light-weight carts are similar for trikes, nanolights, gyroplanes and some powered hang gliders. The wing primary load structure is aluminium tubing plus a lot of hardware fittings forming triangulated structures that are supported by secondary triangulated structures of aluminium tube plus stainless steel rigging cables. The components of the rather complex sail structure are generally cut from polyester materials and sewn together. Shaped battens contained in chord-wise sail pockets provide the aerofoil shape. The sail is only tightly attached to the aluminium frame along its leading edges and wing tips, leaving the trailing edge and much of the rear section of the sail free to flex and twist under load, altering the aerodynamic forces generated by the left and right halves of the wing. In flight the aircraft's centre of gravity (cg) is normally located vertically below the carriage hang point and horizontally near the propeller's extended line of thrust. There is no tailplane and there are no control surfaces like rudders or elevators. Aircraft speed is controlled by rotating the wing, in the pitching plane, about the pitch-and-roll joint thereby altering the wing angle of incidence. To increase speed – for the same power setting – back pressure is held on the control bar (i.e. seemingly pulling the suspended load toward the bar) to reduce wing incidence and thus the angle of attack (aoa). To reduce speed – for the same power setting – forward pressure is applied to the control bar (i.e. seemingly pushing the suspended load away from the bar) to increase incidence and thus aoa. These control bar movements shift the cg fore or aft in relation to the vertical line of the centre of pressure — hence the 'weight-shift' term. The throttle controls climb and descent. In cruising flight the wing centre of pressure is vertically coincident with the aircraft centre of gravity. At slower speeds (higher aoa) the wing cp is vertically aft of the aircraft cg creating a nose-down pitching moment. At higher speeds (lower aoa) the wing cp is vertically forward of the aircraft cg creating a nose-up pitching moment. Aircraft movement in the lateral plane (rolling and subsequently turning) is initiated by the pilot applying sideways pressure on the A-frame control bar — which is fixed relative to the wing. As perhaps 80% or more of the total aircraft mass is represented by the carriage and its occupants and that mass suspended below the wing has considerable inertia then, rather than moving the carriage sideways left or right the control bar movement rotates the wing about the hang point. Consequently the aircraft starts to bank while, at the same time, the action effectively shifts the cg in relation to the wing aerodynamic centre (hence 'weight-shift'). The aoa has to be increased at the same time by forward pressure on the control bar, providing the centripetal force for the turn. The only other flight control is the throttle. As there is no control for rotation about the normal axis, weight-shift aircraft are sometimes referred to as 'two-axis' aircraft. A trike is limited in manoeuvrability; pitch angles of 45° and bank angles of 60° are the recommended maximums; otherwise the usual physics apply for turning, climbing and descending. The hang glider and powered hang glider The very light powered hang glider (PHG) system is similar; the main difference — apart from fewer, and lighter, hardware fittings — is the lack of a carriage or cart. the hang gliders have weight-shift control (i.e. body shift) by the pilot moving their body fore-and-aft or sideways relative to a simple, fixed, triangular control bar and frame system rigidly attached to the wing. The pilot's harness is directly attached at a hang-point on the tubular metal wing keel structure. In addition to the trikes the Hang Gliding Federation of Australia [HGFA] administers a class of powered hang gliders that have an empty weight under 70 kg plus another class, sometimes classified as nanolights, that have a maximum take-off weight under 300 kg. PHGs employ a specialised 10–20 hp paramotor, fuel tank and propeller cage rigidly attached to a light frame within a harness suspended from the hang-point. The pilot is harnessed to the frame in a standing/running position for foot-launching of the aircraft and in a sitting position when airborne. Some PHG, particularly two-place aircraft, may use a light three-wheel cart. The cart relieves much of the physical loads on the pilot when launching and simplifies take-off and landing when carrying a passenger. The non-powered hang glider (HG) system is much the same as the PHG without an engine. Generally the aircraft must be foot-launched. After launching the pilot is either in a seated position or a prone face-down position. For descriptions of some currently available hang gliders, nanolights and trikes see the AirBorne Australia website. 3.12.2 Powered parachutes The ram-air parachute wing The parachute wings used by sport parachutists, paragliders [PG], powered paragliders [PPG] and powered parachutes [PPC] function quite differently to the traditional circular, umbrella-shaped emergency 'chutes. The latter are 'descending only' parachutes, dimensioned so that the drag of the parachute canopy counters much of the weight of the load, limiting the rate of fall to a terminal velocity around 18–25 km/h (5 to 7 m/sec). During the descent the parachute's path acquires a horizontal component as it drifts with the wind, though an encounter with a vertical gust will increase or decrease the rate of descent. The load might be a person who has 'bailed-out' of an aircraft, or even an aircraft plus the occupants — if it's part of a rocket-deployed emergency recovery system. Parachute wings or parawings, on the other hand, also generate lift, allowing the person or aircraft to glide with a fairly low rate of sink and thus to ascend with any parcel of rising air that has an ascent rate exceeding the aircraft's sink rate. Typical L/D ratios for unpowered paragliders are around 8:1. As the 'chute and harness weigh less than 20 kg, L/D much depends on the weight of the pilot and the selection of wing size. Parawings are steerable and provide that high degree of manoeuvrability demonstrated by skydivers and paragliders; and they can be flared for a soft landing. The parawing is generally rectangular in shape; higher aspect ratio elliptical wings provide better performance but are not as stable as a low aspect ratio rectangular wing. Parachute wing construction When the open end of a closed tube is aligned with, and exposed to, a continually moving airflow, the flow within the tube is halted and the rather small amount of pressure energy (see 'stagnation pressure') needed to halt the airflow within the tube is additional to the ambient atmospheric air pressure. This is the basis of the 'ram-air' parachute wing used in the sport parachutes, paragliders and powered parachutes. The design of the skydiving parachutes is a little different from the others as the system must cope with high shock-loads generated as it opens to arrest a free-falling body and the aspect ratio is very low, perhaps less than 2:1 to 2.5:1, to facilitate their very close canopy formation descents. Ram-air wings are formed from a low-porosity material, such as rip-stop nylon, and consist of an upper and a lower fabric surface separated by fabric load-bearing ribs; thus creating a number of individual wing cells open to the airflow at the leading edge and sealed at the trailing edge. The rib fabric is cut in an aerofoil shape (i.e. a parafoil) with interconnecting cross-ports cut into them, so maintaining an equal pressure distribution across a group of cells. In flight, although fabric permeability has a slight effect, the ram-air pressure within the cells is near the stagnation pressure — the highest — and is enough to form the semi-rigid wing shape (a cambered upper surface and a flatter under-surface) that generates lift, providing the gliding/soaring flight ability and the manoeuvrability of parachute wings — as long as the stagnation pressure holds. Once established, the higher stagnation pressure is inside the mouth opening and there is airflow into the cells, then back out over both the upper and lower surfaces. The better designs of parawings have smoother flow. The suspension lines are dimensioned to form the wing into an anhedral arc in flight, thus a PPC usually has a fairly low effective aspect ratio (around 4), but the arc adds to the system's pendular stability because the lift vector at most cell positions will have a lateral component. Turning is accomplished by increasing drag on one side of the wing — by pushing foot pedals or steering bars or pulling steering toggles — which in turn pull down on the brake lines attached to the wing trailing edge. This is supplemented by weight-shift — the pilot leaning. The deflection acts like fully lowering a flap increasing drag on that side and the aircraft yaws and turns. The greater the deflection, the steeper the turn — and the greater the height loss, unless power is increased. Braking both wings simultaneously and reducing power will flare the aircraft for landing (the increased drag slows the wing, the cart swings forward and up a little before touching down); excessive braking may stall the wing. Sport parachutes need fine relative speed, direction and descent adjustment systems for canopy formation manoeuvring. Parawings are used in paragliders, powered paragliders and the powered parachutes described next. The powered parachute aircraft A powered parachute aircraft [PPC] is a two-part system consisting of a cart for one or two occupants with engine and propeller plus the parawing and suspension lines. About 80–90% of the total system drag is contributed by the wing. PPC with rectangular wings have a low L/D — between 3 to 5, but L/D is greater for elliptically shaped wings. PPCs normally cruise at only one aoa and airspeed — around 30–35 knots, although the aoa of some wings can be trimmed in flight to change aoa a little. The aoa of some wings can be changed by shifting weight fore or aft, and maintaining that pilot/passenger position — much the same as altering the trim state of a three-axis very light aircraft by the pilot leaning forward or back. All parawings are capable of stalling (the cells lose their pressure differential and the upper wing surface collapses) if badly mishandled, or if flown in turbulence greater than 'low'. The engine, pilot and passenger are usually accommodated (side-by-side or tandem) in a tricycle undercarriage vehicle — similar to the trike — and often with the parachute lines being led into four attachment points — two forward for the leading edge lines and two aft for the trailing edge lines. The cg is low on the cart, the thrust line is above it and the line of drag is very high. Although it is a two-part system, the two parts act as a whole provided the state of trim is maintained. If power is increased above cruise power, the thrust will initially push the cart forward of the wing — increasing pitch — and the PPC will climb at the designed speed. Rate of climb is dependent on throttle opening and all-up weight. Similarly, if power is decreased, the pitch will decrease and the PPC will descend. In normal cruise, climb and descent, the wing automatically adjusts to the aoa. Pendular stability For pitch and roll stability, the PPC relies on the natural pendular stability provided by the long vertical separation between the aerodynamic centre of the wing and the cg. As a dry nylon wing and suspension lines etc probably weighs less than 20 kg the cg of the two-part system will be within the cart. The wing acts as the suspension point for the 'weight' (the cart and crew) of the pendulum and the suspension lines act as the pendulum rod. Any turbulence will tend to move the wing further than the cart, because of the cart's much higher inertia, and the pendular action quickly restores the normal state after the perturbation — although the normal state is probably a slight gentle oscillation of the cart because of its freedom to swing longitudinally and laterally. In smooth air the PPC can generally be flown 'hands-off'. A gust from the front has the effect of moving the wing back, in relation to the cart. This will temporarily increase aoa and thus lift, because V² is maintained, and the aircraft will rise a little until the cart swings back under the wing and aoa is returned to normal. A gust from the rear has the effect of moving the wing forward, and decreasing aoa and thus lift. The aircraft will sink a little, until the cart swings forward and aoa is returned to normal. Pendular stability is dynamic, so there will be a few oscillations of rising/sinking after such disturbances. Gusts with a vertical component will affect aoa and wing-loading as with three-axis aircraft. In addition to atmospheric disturbances, transient changes in attitude, aoa and airspeed can be induced by over-controlling — fast throttle changes, radical control inputs and fast weight-shifting. The wing will usually — depending on torque at varying rpm settings — turn into the relative airflow and take the cart with it. This can be a problem in the take-off or landing roll if not conducted directly into wind, or if conducted in turbulent conditions. For more PPC information see Aerochute International. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.11.1 Maximum take-off weight Regulatory weight limits All eight recreational aviation exemption Civil Aviation Orders [CAOs] specify a limiting take-off weight or, in a few cases, a limiting empty weight. Take-off weight is the total weight of the aeroplane when it begins to taxi before taking off. The maximum allowed take-off weight [MTOW] has a number of connotations. The first is the class regulatory limit (usually 600 kg but it could be less and up to 850 kg for sailplanes) set by the CASA for recreational aeroplane operations and currently defined in the exemption orders. Those CAOs allow an individual aeroplane to be registered within a class, defined by one particular CAO sub-category, for operation not above a specified take-off weight. In addition there may be a maximum stalling speed or a maximum allowed wing loading specified in those orders. The second connotation is the structural design weight limit which is the maximum all-up take-off weight permitted by the aircraft designer, for structural safety or aircraft stability and control reasons. An aeroplane which, by design, is capable of operating safely at a greater weight than the class regulatory limit may still be able to be registered with a Recreational Aviation Administration Organisation [RAAO], provided the pilot does not operate the aeroplane at an all-up weight that exceeds the class regulatory limit — including the consequent stall speed — defined by the relevant CAO. Many small, light, composite aircraft are imported from Europe where the European Union certification standard for very light aircraft is CS-VLA (formerly JAR-VLA) with a class regulatory limit of 750 kg. These modern technology aircraft have a comparatively low empty weight and potentially high fuel capacity, so it is quite feasible to operate them as two-place 600 kg aeroplanes — provided the combined weight of the occupants is not excessive. There are other older design, two-place, light aircraft where the structural design weight limit is higher than the class regulatory limit. It may be that an RAAO might accept such an aeroplane, but these are required to carry a cockpit placard stating that the MTOW does not exceed 600 kg — or whatever the class regulatory limit might be. Because they have a comparatively high empty weight they must be operated as a single-seat aircraft so permanent removal of the passenger seat, seatbelt, passenger-side controls etc would be required to ensure operation only as a single-place aeroplane. In the type approval process, an aircraft might be assessed by a National Airworthiness Authority [NAA] to determine that the structural design weight limit is considered safe. Subsequently, the third connotation — a maximum total weight authorised [MTWA] — may apply. That MTWA may be less than the structural design weight limit and may be less than the class regulatory limit. The situation is further complicated when overseas factory-built aircraft are imported into Australia for registration with an RAAO. An example is the European countries who certify their aircraft to an European ultralight standard of 450 kg or 472.5 kg (the 22.5 kg is the addition for a parachute recovery system). If imported into Australia and registered with an RAAO, that organisation has no choice but to limit the aircraft to 450 kg or 472.5 kg MTOW even though the class regulatory limit might be 600 kg. However, if the manufacturer certifies them to another standard at a greater weight — providing that certification is accepted by a certifying body in a country that is an ICAO signatory — then an Australian RAAO can accept that higher weight, but only up to our regulatory cut-off point. Australia is an ICAO signatory and the CASA is the NAA and a certifying body. Structural design weight limit From a flight operation and safety viewpoint, the most important MTOW is the structural design weight limit, which may be less than, or greater than, the MTOW allowed under the relevant CAO. The distribution of that weight — the aircraft balance — is equally important. The structural design weight limit is related to the category of operation and the flight envelope. In the 'normal' category — applicable to all ultralights except light sport aircraft [LSA] — the structure, particularly the wing, is required to cope with minimum structural load factors of +3.8g to –1.5g. Thus, the wings of a non-aerobatic aircraft with a certificated MTOW of 600 kg is required to cater for a design limit load of 600 × 3.8 = 2280 kg plus the 50% safety factor for the ultimate load = 3420 kg. No matter which CAO class regulatory limit recreational aircraft are generically permitted to operate at, no aircraft may fly legally above the RAAO accepted MTOW for that particular aircraft type, which may not be as much as the class regulatory limit or the structural design weight limit. Payload Bear in mind that these limits relate to the structural strength of a new aircraft — and structures lose strength as they age; maybe more so if they are a very lightweight structure with little fail-safe provision. However, as aircraft age they also suffer from service weight pickup. They tend to put on weight through modifications, additional instruments or avionics, larger fuel tanks, heavier tyres and accumulation of paint and dirt — inside and outside — all of which reduce payload* capability and make it rather easy to unwittingly exceed MTOW. *In the sport and recreational aircraft context the 'payload' term encompasses the weight of the pilot, passenger(s), baggage and usable fuel. In the general aviation field, most of the privately-owned recreational tourers are single-engine, fixed-undercarriage, four-seat aircraft, like the Piper Warrior or the Cessna 172. Generally these aircraft have a MTOW around 1150 kg — comprising an empty mass which is about 55% of MTOW and a fuel capacity about 15% of MTOW; consequently, 30% of MTOW (around 345 kg) is available for carriage of the pilot, passengers and baggage. Most two-seat light and ultralight aircraft do not have a high payload capability; consequently a full fuel load — which weighs about 0.71 kg/litre — and just an average 80 kg pilot and passenger might constitute, or exceed, the maximum payload. A most important part of pre-flight planning is to ascertain the total weight of the pilot, passenger/s, baggage, tools and other cockpit gear plus fuel. It is also advisable to re-weigh the empty aircraft occasionally to re-establish the empty weight and the cg position when empty. Exceeding MTOW has consequences that increase exponentially with the excess weight: reduced structural load safety factor reduced acceleration, higher take-off speed and longer take-off distance reduced rate and angle of climb reduced cruising speed and range higher stalling speed and reduced manoeuvrability higher landing speed and extended landing distance or maybe the aircraft won't even leave the ground on take-off — which can be a bit expensive if you end up in the fence at the end of the strip. It is much more dangerous if it does get airborne but you trip over the boundary fence (see ground effect) — or if you can't establish a climb rate greater than the vertical velocity of down-flowing air at the end of the runway. If MTOW is exceeded and the cg location is outside its limits, then very dangerous longitudinal stability conditions are introduced. 3.11.2 Balance: containing cg position within limits Balance refers to the location of the cg along the longitudinal axis. Location of the cg across the lateral axis is important, but the design of practically all aircraft is such that the empty weight is generally symmetrical about the longitudinal centreline. However, the location of the cg along the longitudinal axis is both variable and critical for stability. Consequently, the cg position must be assessed by the pilot before every take-off — even if the total weight is well below design maximum safe operating weight. The lateral and longitudinal position of the cg on any flight will vary according to the weight in the pilot and passenger seats, the amount of fuel in the tank(s), the placement of any baggage and other gear, and also the weight and location of modifications and additional installed equipment since the last cg position check. (The load must be properly secured and small objects properly stowed. The last thing you need is a heavy object banging around the cockpit in turbulent conditions and damaging the canopy or something equally vital — like your head. Anything loose in the cockpit/fuselage has potential to jam the control circuits or to move rearward in the fuselage during take-off acceleration or while climbing, thus adversely affecting the cg position.) Centre of gravity range For safe aircraft operation, there must be calculated limits to the forward (nose-heavy) and the aft (tail-heavy) cg position. Those limits — measured from a datum — are specified by the manufacturer or by the amateur designer. (The datum is an imaginary vertical plane through the fuselage, possibly located at the engine firewall, the wing leading edge or perhaps the back of the spinner.) If the cg is situated between the fore and aft limits, the aircraft should have positive static longitudinal stability. Care should be taken when flying amateur-designed aircraft, as the cg range for that aircraft may not be within practical safe limits, making the aircraft dangerously unstable in some conditions. In the 'Aerofoils and wings' module it was stated that the wing aerodynamic centre [ac] is situated near 25% mean aerodynamic chord [MAC]. For longitudinal stability in light aircraft the most forward position of the cg allowable might be about 15% MAC and the most aft position about 35% MAC, basically 10% either side of the wing ac; or perhaps the aircraft neutral point. Forward cg limit — nose-heavy The forward cg limit is determined by the elevator's ability to flare the aircraft at low speed when landing in ground effect; i.e. the least forward cg position where full up-elevator will obtain sufficient moment arm to rotate to the stall aoa, without requiring the pilot to exert an excessive pull on the control column. The forward position is constrained because the further forward it is, the more download the horizontal stabiliser/elevator is required to produce to balance it. Consequently, the tailplane must fly at a greater negative aoa — thus decreasing total aircraft lift — and the wing must then fly at a greater aoa to counter the loss. This results in more drag from the wing and the tailplane and, consequently, reduced performance. The pitching moment characteristics of the wing must also be considered. If a nosewheel undercarriage aircraft is landed in a nose-heavy condition, the possibility of touching down nosewheel first — wheelbarrowing — is greatly exacerbated; a slowing aircraft, pivoting on the nosewheel, is in a grossly unstable condition. The possibility of an extreme ground loop, with consequent aircraft damage, is high. Also touching down nosewheel first can result in a bounce that is difficult to control and may end up wiping off the nosewheel gear and overturning the aircraft. If the c.g. is too far ahead, the aeroplane will continue to be stable but it could be so nose-heavy that it cannot be brought into a landing aoa, that is, it would be difficult to slow it down to landing speed. Aft cg limit — tail-heavy The aft limit is determined by the amount of reduction in the length of the horizontal stabiliser moment arm — which decreases the effectiveness of the moment — and the increase in the nose-up pitching moment of the cg/ac couple, because of the cg distance behind the ac. It is the elevator authority available at low speed that determines the aft cg limit. A cg outside the aft limit will decrease or remove longitudinal stability, and the ability to recover from stalls and spins and may itself lead to a departure stall (i.e. a stall shortly after starting to climb out from the airfield with the engine at maximum power) because there is insufficient elevator authority to lower the nose; even with the pilot applying maximum forward pressure. A go-around with the cg near the aft limit — with flaps extended, full power, and nose-up landing trim still applied — can be particularly dangerous for the unwary pilot. An aircraft does not have to be near MTOW for the fore or aft cg limits to be breached, as can be seen in weight/cg position limitations. The cg position will change as fuel is consumed. Actually, the pilot of a light aircraft can vary the cg position just by leaning forward or backward in the seat! The following is an extract from an RA-Aus incident report: "The aircraft, with instructor and student on board, was returning to the airfield when a pitch-down occurred. (Unknown to them, the elevator control horn assembly had failed.) Control stick and trim inputs failed to correct the situation, but a reduction in power did have a correcting influence, though not enough to regain level flight. A satisfactory flight condition was achieved by the pilots pushing their bodies back as far as possible and hanging their arms rearward. A successful landing at the airfield was accomplished." Mean aerodynamic chord The cg location can be expressed as a percentage of the mean aerodynamic chord [MAC], which is particularly useful for designer/builders. For a rectangular wing of constant aerofoil section dimensions, MAC is just the chord. For a symmetrically tapered wing, it is the average of the root chord and the tip chord. Further information is in 'ascertaining MAC graphically'. The position of the fore and aft cg limits is measured as a percentage of MAC, from the MAC leading edge. Usually for a single- or two-seat aircraft, the most forward position would be aft of 15% MAC and the most aft position would be forward of 30–35% MAC. Thus, the allowable cg range in a light aircraft shouldn't exceed 20% MAC. The linear distance between the fore and aft limits is perhaps 15 to 20 cm. Weight/cg position limitations To demonstrate how the weight and balance limits for a particular aircraft may vary according to the planned flight operation, I have selected a four-seat aircraft that is certificated for operation in three certification categories — normal, utility and acrobatic. The following data is extracted from the aircraft flight manual. The maximum take-off weight (in pounds) for operations in each category are: normal 2335 lb, utility 2137 lb and acrobatic 1940 lb. The fore and aft cg limits are measured in inches from the datum and also shown as a percentage of MAC. The maximum number of persons on board [POB] allowed for each condition is shown. Category Max. weight (pounds) Fwd limit (inches) % MAC Aft limit (inches) % MAC POB Normal 2335 98.19 27 103.58 36 4 Normal 1960 93.07 18.5 103.58 36 2 Utility 2137 95.47 22.5 101.77 33 3 Utility 1960 93.07 18.5 101.77 33 2 Acrobatic 1940 93.07 18.5 97.58 26 2 The table data is summarised below in graphical form, depicting the weight/cg envelope. The vertical axis depicts weight in pounds and the horizontal axis the stations in inches from the datum. The section outlined in blue is for normal operations with a +3.8g limit load factor, the green outline is for utility operations (training, spinning) with +4.4g limit and the red area is for acrobatic category operations with a +6.0g limit. To determine the fore and aft cg limits, first ascertain the weight position on the vertical scale and read across within the appropriate category. For example, with weight 2180 lb in the normal category, the forward cg limit is at 96 inches from the datum and the aft is at 103.58 inches. Note the very restricted cg and MTOW range for aerobatics — 4.51 inches (11.5 cm) or 7.5% MAC — and the requirement for the forward limit to start at 18.5% MAC, the most forward position. On the other hand, when the aircraft is at maximum normal category weight the cg range is only 5.39 inches (13.5 cm) — but now the cg range is required to be at the other end of the scale, between 27% and 36% MAC. The only occasion when the aircraft balance can be anywhere in the specified range of 18.5–36% MAC (10.51 inches or 27 cm) is when the aircraft is operating in the normal category at a weight less than 2000 lb. The area sliced off the top left corner is fairly representative of most weight/cg limitation envelopes for medium to higher-performance light aircraft. Ascertaining mean aerodynamic chord graphically For a rectangular wing of constant aerofoil dimensions and constant chord, the MAC is just the chord. For a symmetrically tapered wing it is the average of the root chord and the tip chord. The diagram below is a representation of the graphical method for calculating the MAC position on such a wing. The method works just as well for more complex wing plan forms. Note that for aerodynamic calculations the aircraft wing includes the area within or above the fuselage and the root chord is always on the fuselage centreline. The position of the wing aerodynamic centre is marked with the red asterisk. 3.11.3 Ballasting Sometimes an aircraft, with a tandem pilot/passenger seating configuration like the Breezy, will require a specified/calculated ballast weight to be strapped in an unoccupied passenger seat, unless the passenger seat is located at the cg position. There are also pusher engine designs that are entirely dependent upon sufficient minimum pilot weight to put them in balance range, so a lightweight pilot may have to sit on a ballast bag. With tandem two-seaters there will be both a minimum and maximum pilot weight for cg range, but that in turn could be influenced by rear seat weight to keep within MTOW. In some cases, the rear seat also has a moment arm and can affect the front seat arm, depending upon rear seat weight. Regulations require that any ballast, baggage or other cargo that is stowed on a passenger seat must not weigh more than 77 kg; the weight should be evenly distributed and positioned so that neither the cargo nor its restraints can interfere with the operation of the aircraft controls. In addition, if fitted with removable dual controls, the control column at the passenger seat should be removed. It is advisable that the cockpits of two-seaters — particularly tandems — but any aircraft that is dependent upon the presence of a minimum and maximum pilot weight, should be clearly placarded with the minimum/maximum seat weights shown in the flight manual. The need for ballasting is not confined to ultralights. The cg position of the four-seat Beech Sundowner is outside the forward limit when the only occupants are two above-average weight people in the front seats, and in such conditions the aircraft has a tendency to wheelbarrow on landing. Flying an unbalanced ultralight — i.e. in a tail-heavy or a nose-heavy condition — even though the cg is not outside the limits, increases pilot fatigue because of the need to maintain a constant heavy pressure on the control column if no trim device, or a limited device, is fitted. 3.11.4 Calculating cg position and moment Unloaded aircraft The longitudinal position of the cg and its moment about a datum are readily calculated. A measuring tape, heavy-duty bathroom scales, plumb bob and a chalk line are needed. The following is the procedure for an empty light nosewheel aircraft. Chalk a straight line on a level surface that is at least the length of the fuselage, then chalk another line perpendicular to that. Roll the aircraft along the longitudinal line until the axles of both main wheels are directly over the cross line. Chalk another short cross line to mark the nosewheel axle position. Mark a position on the longitudinal line that is directly below the front or back end of the spinner thus providing a datum. Measure the longitudinal distance (the nosewheel moment arm) from the datum to the nosewheel axle line and the distance (the mainwheel moment arm) from the datum to the main wheels axle line. Place the scales under the nosewheel, block up the mainwheels so that the aircraft remains level and note the weight. Then place the scales under one of the mainwheels and block up the other main plus the nosewheel. Note that weight. Repeat for the other mainwheel. Add the weight on the nosewheel to arrive at the aircraft empty weight (or perhaps its weight with full fuel). Multiply the nosewheel weight by its arm to get the nosewheel moment and the added mainwheel weights by the axle arm to get the mainwheel moment. Add the two together to arrive at the total empty aircraft moment. The cg location from the datum equals the empty aircraft moment divided by the total aircraft weight. For example: Nosewheel weight = 80 kg and arm = 0.5 m Thus nosewheel moment = 40 Mainwheel weight = 2×160 kg and arm = 2.5 m Thus mainwheels moment = 800 Empty aircraft weight = 80+160+160 = 400 kg Empty aircraft moment about the datum = 40 + 800 = 840 Cg location when empty = 840/400 = 2.1 m from the datum Loaded aircraft The cg location with pilot/passenger aboard can be calculated if a point about 20 cm forward of the seat back (being the approximate centre of mass position of a seated occupant) is marked on the longitudinal chalk line; the distance from the datum to that point is the front seat(s) moment arm. The front seat(s) moment is the occupant(s) weight multiplied by the arm, and the new cg location is the empty aircraft moment plus the front seat moment divided by the empty aircraft weight plus occupant weight. For example: Side-by-side front seats arm = 2.3 m Occupants weight = 150 kg Thus front seats moment = 345 Empty aircraft weight = 400 kg Empty aircraft moment = 840 Total aircraft weight = 550 kg Total aircraft moment = 1145 cg location = 1345/550 = 2.08 m from the datum Similar calculations can be made to include fuel weight and baggage distribution and weight. Aircraft or kit manufacturers should provide data defining a datum together with the associated arms for the pilot/passenger seats, fuel tanks and baggage compartments, plus the fore and aft cg limits expressed as a distance from the datum. With such information the pilot can calculate the loaded cg position using the measured weights of occupants, fuel and baggage. The aircraft manufacturer should provide a loading chart to facilitate calculations. Of course the manufacturer's chart is useless (and you may make the aircraft dangerously unstable) if you stuff baggage and equipment into any available space outside the designated and designed baggage compartment. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.10.1 Control in pitch In section 6.3 we learned that movement of the elevators provides a pitching moment about the lateral axis, that initiates a change in the aoa. Once the new aoa is established, then — provided the elevators are held in that deflected position by pilot pressure on the control column or a trim device — the pitch moment returns to zero and the aircraft maintains that aoa (and there is a direct relationship between aoa and IAS). In the manoeuvring forces module we established that aoa and power combinations provide (a) increased speed or climb and (b) decreased speed or descent — or varying degrees of either. Thus, control in pitch (i.e. of aoa) combined with throttle control allows an aircraft to take-off, climb, cruise at various speeds, descend and land. However, control in pitch involves more than initiating a discrete pitching moment to effect an aoa change and subsequent attitude change. In section 1.10 we found that to sustain a turn, an additional force must be continuously applied towards the centre of the curve or arc — the centripetal force. This is achieved by an increased aoa — greater than the normal for a particular straight and level airspeed — held with control column back-pressure. The increased aoa provides the centripetal force, and that force keeps the aircraft constantly pitching 'up' in the longitudinal plane, into the direction of turn. 3.10.2 Control in yaw Aircraft perform much better if their longitudinal axis is accurately aligned, in plan view, with the flight path; i.e. with the relative airflow. Also, an aircraft flying with a constant angle of slip is part way to ending up in an unusual flight attitude. Thus, the primary task of the fin and rudder combination is to act as a stability and trim device so that the directional stability system will restore flight to the proper state of zero sideslip angle. If the rudder has no trim tab, as would be the case with most light aircraft, then the pilot would have to keep the aircraft in trim by applying pressure to one rudder pedal. This pressure will vary with airspeed and the sideslip angle. This trimming task also includes the use of rudder to overcome adverse yaw when initiating a turn, and to keep the turn balanced or 'coordinated'. The very simple flight instrument provided to indicate slip — or skid in a non-coordinated turn — is the balance ball. A metal ball is enclosed in a short, transparent, slightly curved tube where movement is somewhat damped by the restriction of the tube. When the aircraft is flying with zero sideslip, the ball will be centred at the bottom of the curve; when the aircraft is slipping into (or skidding out of) the turn, the inertial forces will move the ball left or right in the direction of the slip. To trim the aircraft, the pilot applies pressure on the rudder pedal on the side to which the ball has moved; i.e. 'steps on the ball'. When the aircraft is slipping, the pilot will also feel those inertial forces apparently pushing his/her weight in the same direction as the ball, hence the expression 'flying by the seat of your pants'. The amount of pilot-induced yaw, at a given airspeed, is dependent on the degree of rudder deflection. If the pilot holds the rudder deflection, the aircraft will continue yawing and sideslipping. But, as the aircraft rotates about the normal axis, the wing on the outside of the rotation must be moving a little faster — and the inner wing a little slower — so there will be a small lift differential. The differing lift moments will raise the outer wing and lower the inner wing and the aircraft will enter a banked turn. A pilot would not initiate a sustained turn by using rudder alone, but there are occasions when it is appropriate and effective to alter the aircraft's heading a few degrees by using just rudder — or perhaps rudder plus a little opposite aileron to stop the bank. Such occasions are when finally aligning the aircraft with the runway centre-line or compensating for small changes in wind direction when landing. We will cover this in the 'Circuit, approach and landing' module. Having said that, there is still an occasion where a rapid 180° change in direction is achieved solely with rudder; see the following. The ultimate use of rudder to yaw the aircraft is demonstrated when the pilot of an aerobatic aircraft executes a 'stall turn' or 'hammerhead'. The former term is perhaps a misnomer, because the wing does not reach the critical aoa during the manoeuvre. The manoeuvre involves pulling the aircraft into a full-power vertical climb then, as the airspeed decays to somewhere near the normal stall speed, applying full rudder so that the aircraft is yawed 180°, about the normal axis, into a vertical descent. The slipstream supplies the energy to the rudder for the turn. The interesting thing is that, although the aircraft is at or below normal stall speed, the wings are nowhere near the critical aoa as, towards the end of the climb, the control column is held forward of the neutral position — and because of the low aoa and low V², the wings are not producing much aerodynamic force*. The length of the vertical 'up-line' depends on the aircraft's power/weight ratio — and thrust power and momentum are reducing quickly. Of course, if the pilot delays too long before applying full rudder, the weight vector will take over and cause the aircraft to slide vertically backwards — a tailslide. *In fact the wing could be at the 'zero lift' aoa where the aerodynamic forces on both sides of the wing are equal and opposite. 3.10.3 Control in roll We discussed how ailerons produce a rolling moment in section 4.10, so what happens when the ailerons are normally deflected, by the pilot moving the control column to the left or right? Initially the aircraft will start to roll, and if the control column is then returned to the neutral position the roll will cease but the bank angle reached will tend to remain. To level the wings, the column has to be moved to the opposite side — then returned to neutral once the wings are again level. Which indicates there always tends to be a sort of 'neutral stability' in the lateral plane. However, that situation doesn't exist because, as we found in section 7.4, other forces come into play when the aircraft is banked — creating sideslip, then yaw and eventually a turn. So the prime reason for introducing a roll is as the first step in turning. However, before going on to the turn, let's just look a little further at the effect of aileron deflection while elevator and rudder are held in the neutral position, when the aircraft's velocity is high; i.e. pure roll. The simplest aerobatic manoeuvre is the aileron roll, which (in aircraft that have been certificated for aerobatics) is usually accomplished by first gaining some extra kinetic energy, applying full power and raising the nose so that it is pointing 20° or so above the horizon with the balance ball centred. The control column is then firmly moved left or right to the full extent of travel and held there, while the elevators and rudder are both held in the neutral position; i.e. the roll is produced solely by the aileron deflection. The aircraft will then continue to roll about its longitudinal axis — because the ailerons being positioned towards the wingtips produce a strong rolling moment about the axis — more or less at a steady rate of roll, until the accumulated extra energy is exhausted. If the column was not moved to its full extent, the aircraft will still roll, but the rate of roll will be slower yet still steady. The time to complete a full 360° roll is very much dependent on the aircraft design and that roll rate dictates how many complete rolls can be produced before the aircraft ends up with the nose pointed well below the horizon — because of the combined effects of slip, yaw and the full 360° inclination of the lift vector. The rate of roll for a competition-class aerobatic aircraft is around 360° per second; for the more mundane aerobatic aircraft, it is around 60° per second, which will probably produce only one full 360° roll. A properly executed roll will result in a continuing 1g load throughout the manoeuvre. 3.10.4 Control in a turn We come then to the question 'how do you turn an aircraft?' Well, you can make the aircraft's nose turn just a few degrees without banking by just applying pressure on the rudder pedal in the direction you want the nose to move and at the same time moving the control column a little sideways in the opposite direction to stop the consequent bank. Applying pressure to the rudder in one direction with opposite aileron is cross-controlling. Normally, this is a very sloppy way to fly but also a habit that can lead to trouble — particularly in low-speed descending turns, such as that made in the approach to landing. Although not directly related to turns, this extract from an RA-Aus incident report illustrates how easy it is to get into difficulties if you don't realise you are cross-controlled at low speeds. "The student had completed two solo circuits and landings without incident. During the third the landing appeared normal, the aircraft touched down without bouncing but then veered left and the left wing lifted. The student applied full power but the aircraft failed to climb normally and appeared to be staggering and slowly orbiting to the left. The aircraft only gained about 40 feet height then gradually descended, striking the ground nose low and left wing low. The student was not injured. It was found that he had maintained full left rudder when he applied full power and was using aileron to counter the yaw — the aircraft basically sideslipped into the ground." The normally recommended way to initiate a level turn (to the left) is to move the control column to the left until the required bank angle is achieved, then return the control column to neutral. At the same time as applying aileron, just sufficient bottom (left) rudder is applied to balance the turn so that there is no slip or skid and the balance ball stays centred. Also, the amount of rudder required increases as airspeed decreases. As the aircraft banks, the lift vector departs from the vertical, so the aoa must be increased sufficiently that the vertical component of lift always exactly balances weight. This means increasing back-pressure on the control column as bank is applied. As aoa increases, induced drag increases. So, to maintain V² throughout the turn, power must be increased. Thus a properly balanced, constant rate and constant-speed turn implies a smoothly coordinated application of aileron, rudder, elevator and power. In some aircraft, particularly slower aircraft with high aspect ratio wings, it is necessary to lead the turn with quite a bit of rudder (because of aileron drag) before adding aileron. In other aircraft it is quite easy to initiate and continue a turn without using rudder at all, but the turn will be uncoordinated — i.e. the balance ball not centred — and such conditions are not desirable. During a banked level turn, the outer wing is moving very slightly faster than the inner wing and will consequently produce more lift; the bank will tend to increase and the turn to wrap-up even though the ailerons are in the neutral position. In order to maintain the required bank angle it is necessary to apply a slight opposite pressure to the control column, which is known as 'holding off bank'. This is relative to level and climbing turns, but different physics apply to descending turns. In a climbing turn, the outer wing has a slightly greater aoa than the inner wing, and thus additional lift. Combined with its slightly faster speed, this reinforces the tendency for the bank angle to increase and the need to hold off bank. The reason for the higher aoa of the outer wing is because of a difference in relative airflow. Imagine an aircraft doing one complete rotation of a continuing climbing turn. Obviously all points on the airframe are going to take the same time to achieve the higher altitude; however, the upward spiral path followed by the outer wingtip must have a larger radius than that followed by the inner, and therefore the path followed by the outer wingtip is not as steep as that followed by the inner. The less steep path of the outer wing (i.e. the relative airflow) means that the aoa of the outer wing will be greater than that of the inner. You might have to think about it a bit! The reverse occurs in a descending turn — the steeper path of the inner wing in the downward spiral means that it will have a larger aoa than the outer wing, which may compensate, or overcompensate, for the faster velocity of the outer wing. In order then to maintain the required bank angle it is necessary to apply an inward pressure to the control column; i.e. in a descending turn the bank must be 'held on'. If the pilot tends to hold off bank in such a turn, there will be an excess of 'bottom' rudder and the aircraft must be skidding. Whenever an aircraft is slipping or skidding, the wing on the side to which the rudder is deflected will stall before the other, with a consequent instantaneous roll in that direction. So the situation we've described — holding off bank in the descending turn with excess bottom rudder — means that should the aircraft inadvertently stall — a cross-controlled stall — it is going to roll further into the bank and enter an incipient spin. Hence the old adage — 'never hold off bank in a gliding turn'. A cross-controlled stall typically occurs in the turn onto final approach for landing. If you must fly cross-controlled when banked, then it is better to fly with an excess of top rudder, as in the sideslip manoeuvre. Thus, if the aircraft should stall, the roll will be in the direction of the upper wing; i.e. towards an upright position. And never apply an excess of bottom rudder in an attempt to tighten any turn, particularly when the airspeed is low for the bank angle employed and/or height is low. This is discussed further in the 'Safety: control loss in turns' module. A breaking turn is a defensive flying manoeuvre, which every pilot should be able to perform rapidly and automatically to avoid collision, particularly in the circuit. It involves very rapid transition, usually into a steep descending turn, but a steep climbing turn may be necessary. A level turn is an unlikely choice, but whatever turn is chosen you must be able to perform it instinctively while your head is continually swivelling to ascertain the location of other aircraft — without falling out of the sky by inadvertently applying back-pressure on the control column and thus exceeding the critical aoa. Before we go on to the sideslip, another very simple aerodynamic demonstration — but only for aerobatic aircraft whose engine and airframe are able to take the loads, and absolutely never for any other aircraft — is the flick roll. The flick or snap roll asymmetric aerobatic manoeuvre is an accelerated stall combining a rapid increase in aoa with a full yaw. It is brought about — when cruising straight and level at a speed less than Va — by pulling the control column firmly back as far as it will go while applying full left or right rudder. The wing in the direction of the applied rudder stalls first and the aircraft flicks into a 360°, roughly horizontal, roll. The roll will continue while the rudder and elevators are held at their limits (and cease when they are returned to the neutral position) but the aerodynamic drag produced by the manoeuvre slows the aircraft quite rapidly and the aircraft will enter a vertical spin if the roll is held too long. The faster the entry speed, the higher the torsional stress on the rear fuselage and empennage, the engine mountings and perhaps even the engine crankshaft; but the slower the speed, the greater the likelihood of immediately entering a spin, and recovery technique is dependent on several variables. The roll is usually a lot snappier in the opposite direction to propeller rotation. Flick rolls may be executed only in aerobatic aircraft designed to withstand the extreme stresses. Such a simple control action, whether or not while turning, demonstrates how easily misuse of rudder can end up in an unusual and dangerous attitude, and where the possibilities increase as speed decreases. And be aware that you don't have to actually push on the rudder pedal, you can easily achieve misuse by inadvertently slipping one foot off the rudder bar at a critical time — the turn onto 'final approach' for example. 3.10.5 Sideslip as a manoeuvre Types of sideslip vary in degree — from inadvertently flying cross-controlled in the cruise (i.e. one wing slightly low and compensating with opposite rudder) to a fully-fledged cross-controlled turn where the aircraft is steeply banked in a descending turn with full opposite rudder applied. All sideslips reflect uncoordinated flight and result in increased drag. Note: in aerodynamic terms, any time it is evident that the aircraft's longitudinal axis is at an angle to its flight path (in plan view) then the aircraft is sideslipping (i.e. its motion has a lateral component), and the angle between the flight path and the axis is the sideslip angle. Aerodynamicists don't generally distinguish between sideslip and 'slip' or 'skid', but many pilots use 'slip' as the general term, 'skid' to describe slipping away from the centre of a turn and 'sideslip' to describe a particular type of height-loss manoeuvre. The sideslipping manoeuvre is only for the pilot who has a very good feel for their aircraft because, among other things, the ASI will most likely be providing a false airspeed indication. High sideslip angles combined with high aoa must be avoided. There seem to be as many definitions of the types of slip as there are exponents of sideslip techniques, but the safe execution of all sideslips requires adequate instruction and continuing practice. Here are some types: The straight or steady-state sideslip approach to landing The helmet and goggles crowd who, very sensibly, like to fly biplanes and other open-cockpit aircraft not equipped with flaps, need a manoeuvre for use on the landing approach to a short strip that enables them to lose height quickly without increasing airspeed and which provides a good view of the landing area. The answer has long been the cross-controlled steady state sideslip; a manoeuvre designed to lose height over a short distance, dumping the potential energy of height by converting it to drag turbulence rather than kinetic energy. Such sideslips may also be a requirement when executing a forced landing, and the same type of slipping approach may also be necessary for those aircraft where, in a normal approach, the pilot's view of the runway is obstructed by the nose. Once established on the approach descent path at the correct airspeed, the aircraft is banked with sufficient opposite (top) rudder applied to stop the directional stability yawing the nose into the relative airflow and thus turning. Slight additional backward pressure on the control column may be needed to keep the nose from dropping too far. The aircraft sideslips in a moderate to steep bank with the fuselage angled across the flight path, giving the pilot a very good view of the landing area. The greatly increased drag, from the exposure of the fuselage side or 'keel' surfaces to the oncoming airflow, enables an increased angle of descent without an increase in the approach airspeed. The execution of a sideslip to a landing varies from aircraft to aircraft and it may not work particularly well where there is a lack of keel surface — an open-frame aircraft like the Breezy, for example. The sink rate is controlled by aileron and power is held constant, usually at idle/low power, and the sideslip must be eased off before the flare and touchdown. When recovering, care must be taken to coordinate relaxation of the back-pressure, leveling of the wings and straightening of the rudder — otherwise the aircraft may do its own thing or stall, particularly in turbulent conditions. The straight sideslip is limited by the maximum rudder authority available; there will be a bank angle beyond which full opposite rudder will not stop the aircraft from turning. Although this manoeuvre usually comes under the proprietorship of the 'stick and rudder' people, the use of the sideslip, by the captain of a Boeing 767, undoubtedly saved the lives of many people in an extraordinary incident that occurred in 1983 when, due to a train of errors — as are most accidents/incidents — an Air Canada 767 ran out of fuel at 41 000 feet. The captain subsequently glided the aircraft to a safe landing on an out-of-service runway, which was being used for a drag racing event at the time. The aircraft was sideslipped through several thousand feet to lose excess height on the approach. For more information about this magnificent demonstration of airmanship (following an execrable demonstration of preflight procedure by many people; keep the old adage in mind — "proper pre-flight procedure precludes poor performance"!) google the phrase 'Gimli glider'. The sideslipping turn Slipping whilst turning is a manoeuvre often used in non-flap equipped aerobatic aircraft where it is desirable to perform a curving landing approach. This is also a useful emergency manoeuvre if it is necessary to increase the sink rate during a turn — such as the turn onto final approach in a forced landing when an overshoot of the landing site is apparent. It is just a sideslip where insufficient top rudder is applied to stop the aircraft turning while slipping. The rate of turn and the rate of sink are controlled by the amount of bank and the amount of rudder but it is an uncoordinated descending turn. Dangerously high descent rates are achieved if the bank angle applied exceeds the full rudder authority. Fishtailing Fishtailing is a series of sideslips where the wings are held level in the approach attitude with (alternating) aileron while the aircraft is repeatedly yawed from side to side by applying alternate rudder; the increased drag increases the sink rate and is possibly used as an emergency measure if overshooting a forced landing. The manoeuvre is generally not recommended, because uncoordinated control use at low levels may lead to dangerous loss of control. Also, excessive alternating rudder reversals may overstress the fin/rear fuselage. Sideslip to a crosswind landing In a sideslip to a crosswind landing, the aircraft is always banked with the into-wind wing down so that the sideslip can be smoothly decreased to a forward slip (below) before the roundout. Most aircraft tend to be slower in the slip, so the nose will need to be a bit lower than that needed to maintain the normal approach speed. A smoothly executed sideslip approach requires much practice, but displays considerable finesse to a ground observer. The forward slip crosswind approach A 'forward' slip is a moderate sideslip application designed only to compensate for crosswind during approach and landing. The slip can be applied throughout the final approach or just in the last stages, and it usually follows a full sideslip approach in crosswind conditions. The into-wind wing is lowered with sufficient bank so that the slip is exactly negating the crosswind drift, while opposite (top) rudder is applied to stop a turn developing and to align the aircraft's longitudinal axis with the flight path — and the runway centreline. If drifting off the path, just add or remove some aileron pressure and, at the same time, add or remove some rudder pressure to maintain direction. An approach speed 2–3 knots above normal is set up, the sink rate (which will be greater than usual because of the inclined lift vector) is controlled by the power setting, the into-wind main landing gear will touch down first and the aircraft is held straight with rudder by pivoting on that one wheel until ground speed has reduced to a safe level. The forward slip is the particularly recommended technique for crosswind landings in high-wing taildragger aircraft. Incidently, a useful technique for a high-wing taildragger in a significant crosswind is to also perform the take-off run on one main wheel. If there is any real difference between the straight sideslip and the forward slip it is just the amount of pressure applied to the controls. In a sideslip, the aileron pressure dictates the angle of descent and the rudder pressure dictates the amount the fuselage is deflected across the flight path. In a forward slip, the aileron pressure is just enough to compensate for the crosswind drift and thus maintain position on the extended runway line, and the rudder pressure just enough to keep the fuselage aligned with both the landing path and the flight path. There is one manoeuvre for certified aerobatic aircraft that demonstrates what might be considered a reversal of all we have stated in this module. This is the 'four-point slow roll' or 'hesitation roll' where the aircraft is rolled through 360° in level flight around a point on the horizon, but the roll is paused for a second or two at each 90° point; i.e. when the wings are first vertical, when the aircraft is upside down, when the wings are again vertical and when the aircraft returns to normal attitude. The roll is started (to the left) with normal aileron and a bit of left rudder, then as the roll progresses through the first 90° top (right) rudder is increasingly applied to negate the yaw, and also to hold the nose up. During the slight pause at the 90° position the aircraft is being held in a nose-up attitude by the rudder whilst the elevators are used to stop the nose wandering to the left or right across the horizon, and the ailerons are neutral. Some lift will be generated by the fuselage having an aoa because the nose is being held up. Then the roll is restarted until, at the 180° position, the aircraft is inverted and the nose is held up by a large forward movement of the control column and the aoa is negative; i.e. the lift is being generated by a reversed aerofoil. And so the roll continues. Of course, all the control movements involve gradual increase/decrease in pressures throughout the sequence. 3.10.6 Spins The 'stall/spin' phenomenon When an aircraft is held in a turn, and near the critical aoa with excess bottom rudder applied (i.e. a cross-controlled skidding turn, which often happens when the pilot tries to 'hurry' the turn with bottom/inside rudder instead of increasing bank) the lower, partly blanketed, wing will be producing less lift than the upper wing. Any tightening of back-pressure on the control column (or any inadvertent back-pressure applied when, for instance, looking over your shoulder; or even any encountered turbulence or wind shear) may take the aoa past the critical angle. The lower wing will drop sharply in an 'uncommanded' roll, and thus become more deeply stalled than the upgoing wing — which may not be stalled or just partly stalled. The high aoa of the lower wing causes greatly increased induced drag, yawing forces in the same direction as the lower wing come into action, the nose swings down and the aircraft enters an incipient (i.e. partly developed) spin condition, where it is about to start autorotation (below). All of this happens quickly, and in some aircraft very quickly indeed. If the aircraft is properly weight-balanced (i.e. weight less than MTOW and cg within the defined fore and aft limits), this is readily countered by quickly easing stick back-pressure to reduce aoa below critical aoa — which immediately restores full control — applying sufficient rudder to stop further yaw, adding full power and rolling the wings level with aileron while keeping the balance ball centred. However, if the slew, exacerbated by the yaw, is allowed to develop past perhaps a 45° movement in azimuth, then there is a stall/spin situation. You sometimes hear about this when an unaware pilot allows such to develop close to the ground — often when performing a climbing turn after departure from an airfield; or in a turn-back to the runway following engine failure after take-off; or when in the descending turn from base leg on to the final approach to landing where, because of illusory ground reference cues, there may be a tendency to increase the rate of turn by applying additional bottom rudder whilst maintaining the bank angle with opposite aileron — 'holding off bank' as mentioned in section 8.4. Ultralights, with their very low wing loading, normally display quite benign stall characteristics when slowly decelerated to stall speed in straight and level flight. But they may exhibit quite nasty behaviour in an accelerated stall or when a stall is initiated during a turn; and such are the usual unintentional stall/spin modes. Under these circumstances the height lost in the incipient spin — the initial entry to autorotation — may be 100 to 400 feet. Thus, an incipient spin condition is highly dangerous when occurring in the circuit pattern or in any other low-level flight situation. See the flick roll box above and read the sections dealing with 'limiting climbing turns during take-off' and 'accelerated stalls'. There may be uncertified light aircraft where characteristics, such as rolling inverted in a relatively mild stall/incipient spin situation, are evident. However, a commercially manufactured aircraft with such characteristics should not receive a CASA Certificate of Type Approval. Autorotation — the fully developed spin With adequate training, an incipient spin is readily anticipated and easy to correct — provided the aircraft weight and balance are within the stated limits. But, if the correction is not done before the nose has swung maybe 90° or so, it may develop into autorotation where the aircraft is descending in a stabilised, usually nose-down, rotation — rolling and yawing in the same direction at a constant airspeed at or slightly above Vs1 — a full-blown spin with each 360° rotation taking only 2–4 seconds in a very light aircraft. The height loss during each rotation — 200 to 400 feet or more, depending on the stall speed and the steepness of the spin — plus the considerable height loss during the pull-out from the recovery dive, is insignificant at a reasonable height but will be critical at lower levels. In a normal turn the aircraft's longitudinal axis is more or less aligned along the flight path — which is the periphery of the turn — the cg moves along the flight path and the inner wing is pointing towards the centre of the turn. But in fully developed autorotation the vertical axis of the spin is located somewhere in the 90° sector between the lateral and longitudinal axes and not so far from the aircraft's cg — perhaps less than two fuselage lengths. Thus, the aircraft is not turning in the normally accepted meaning of the word; it is 'spinning' around that vertical axis, while it's also rolling and yawing about the aircraft's cg; and also pitching somewhat. The lower wing is more deeply stalled than the higher producing less lift but, being on the back slope of the CL curve, more induced drag so providing the asymmetrical yawing and rolling moments. Those aerodynamic forces produced by the wings drive the spin while the resistance of the rear fuselage and empennage reaches a point where it prevents the yaw from developing further; the aircraft's inertia resists change in angular momentum so producing the stable autorotation condition. Usually the structural loads are only a little above normal during autorotation. In a steep spin, the nose is pitched down perhaps 50–60°, the aoa of the lower wing is 20–30°, there is a fair bit of bank and the roll motion dominates. The spin axis will be perhaps somewhere near one or two fuselage lengths forward (more or less) of the aircraft's cg — further away for a steeper spin. The cg will be following a helical flight path. In a flat spin, the nose is pitched down perhaps 10–20°, with an aoa around 60–70° due to the high vertical component of the relative airflow. With very high induced drag and little bank, the angular rotation winds up and yaw motion dominates. The spin axis will be much closer to the aircraft's cg, perhaps even within the airframe, and particularly so if the cg is in an aft position. The closer the spin axis is to the cg, the harder it is to break out of the spin. If the axis coincided with the cg, break-out would be impossible — unless the aircraft was equipped with a ballistic parachute recovery system. Most very light tractor-engined aircraft spin steeply to moderately steeply, so spin recovery early in autorotation is usually — but not always — straightforward: close the throttle, ailerons to neutral, stop the yaw (by applying full rudder opposite to the rotation direction apparent through the windscreen or shown by the turn indicator (not the balance ball/needle), then unstall the wings to stop the spinning (generally by applying full forward stick rather than just moving it to or past the neutral position until the spin stops). Control movements must be carefully sequenced and positive. The aircraft will be in a steep descent when the spin has ceased; the aerodynamic loads during the subsequent pull-out from the descent may lead to an accelerated stall if the aircraft is nearing the surface and the pilot applies extreme back-pressure. The height loss just during the pull-out stage may exceed 400 feet, so that the total loss of height during spin entry and recovery could easily exceed 1000 feet. The problem for a pilot who is conscious of the need to avoid stall conditions when in the circuit by always maintaining a safe speed near the ground, and has had ample training in stall and incipient spin recognition and recovery, occurs when a spin is inadvertently induced at altitude. If that pilot has never previously encountered full autorotation then the disorientation associated with the first experience can be frightening. The pilot may also experience a ground rush illusion where the surface features rapidly spread out to fill the entire field of view and the ground appears to rapidly rise; the reaction is to freeze or to pull back on the control column, which just ensures that the aircraft is held in the stalled condition even though there may be ample height available to recover. The photo at left was taken about 1949 and shows what happened when a student pilot got himself into a spin, evidently retained back-pressure on the control column and allowed the Tiger Moth to spin all the way to the ground from above 2000 feet. The spin developed into a flat spin, with relatively low vertical and horizontal speed, enabling the pilot to walk away with minor injuries. Also the Tiger Moth had a tough steel tubing fuselage frame, which absorbed much of the impact energy. You can see that the fuselage aft of the engine compartment firewall seems practically undamaged. The pilot of any aircraft will not be exposed to the risk of an unintentional stall/spin if they always remain situationally aware, maintain an appropriate energy balance, does not indulge in very low-level manoeuvring and, above all, flies the aircraft. Don't practice stalling below 3000 feet agl; and remember spins result from a loss of lateral and directional stability at the critical aoa, and the only way to get into a spin is to first exceed the critical aoa. Also, sufficient forward stick movement will immediately decrease aoa below the stall angle and restore full control in any stall or near-stall condition; but not in autorotation where opposite rudder and full forward control column movement is necessary because (1) the aoa developed will be well past the critical aoa and (2) the control surfaces will not be as effective as usual — the fin and rudder could be screened by the tailplane and thus in a low energy, turbulent airstream. Intentional spinning Stall/spin events occurring at a safe altitude are insignificant; they may become nasty accidents when they occur at lower levels. But intentional spins can be fun in an aerobatic aircraft that has been certificated for intentional spinning and, given sufficient height, quite easy to recover from — provided the cg position is within the forward and aft limits. The challenge is in aiming for a precise number of turns, half-turns or quarter-turns and the exact direction the aircraft will be heading at recovery. Spin characteristics are very complex and vary greatly between aircraft. Generally the intentional spin is induced from level flight by closing the throttle, bringing the aircraft to the point of stall in a nose-up attitude, holding ailerons in the neutral position then applying full rudder in the direction you want the aircraft to rotate and, at the same time, pulling the stick right back. Hold the neutral aileron, full rudder and back stick. The reason for the excessive control movements is to ensure a swift and definite entry into autorotation. The higher the nose is held above the horizon at the point of stall the more violent will be the spin entry. Aircraft that tend to spin with the nose pitched well down will recover more quickly than aircraft where the spin attitude is relatively flat. However, if allowed to continue past two or three full turns, then centrifugal forces become well established — which tend to make all parts of the aircraft rotate in the same horizontal plane. Then, a nose-down spin may turn into a flat spin, which will then speed up rotationally, the rate of descent decreases, spin radius decreases and break-out will take longer, or may not be possible because it may be impossible to lower the nose. The spin axis may be very close to the pilot which would be very disconcerting. Recovery control forces required usually increase as the spin winds up; also, after initiating recovery action, the spin may increase a little before the action takes effect. Engine power — and its associated effects — also tends to flatten the spin. The flatter the spin, the closer the spin axis is to the cg and the greater the aoa, maybe 75° or more! Also, at such angles, the rudder may be completely blanketed by the fuselage/tailplane, making that control quite ineffective. Structural stresses increase as the spin progresses. A flat spin might be induced if, at the point of stall, full opposite aileron is applied with full rudder. If an aircraft stalls when inverted, it may enter an inverted spin if the control column position was held well ahead of neutral at the stall. It only happens during aerobatic routines — such as a poorly executed entry into a half-roll off the top of a loop, or messing up a stall turn. The recovery from an inverted spin involves correcting the yaw and increasing stick back-pressure until rotation ceases, then rolling level when speed has increased sufficiently; but the great danger in an inverted spin is pilot disorientation. One thing is certain — NEVER, NEVER intentionally spin an aircraft that has not been through the complete spin certification process; they may be incapable of recovery from fully developed autorotation, or the recovery attempt may result in a violent manoeuvre that overloads the airframe. Spin restrictions are not confined to non-aerobatic aircraft; for example, intentional spins were prohibited in the Seafire 47 and Sea Fury, very fast naval fighters of the late 1940s early 1950s, because of the time to recover (if recovery was possible) and the consequent extreme height loss. Spin recovery confidence building Developed spin recovery training is not included in the RA-Aus Pilot Certificate syllabus or the General Aviation Private Pilot Licence syllabus, but stall and incipient spin awareness and recovery are normal parts of the syllabi. A spin is usually classified as an aerobatic manoeuvre and, as all RA-Aus registered aircraft are prohibited from such manoeuvres, they shall not be allowed to enter an intentional developed spin. More to the point, no ultralight (and rather few light non-aerobatic aircraft) has ever been through the complete flight test schedule for spin recovery. However, gaining some experience and confidence in recovery from full autorotation can be readily and cheaply obtained by practicing a half-dozen spin recoveries with an instructor in a two-seat glider or powered aerobatic GA aircraft. It is probably better experience in the glider, as you are also exposed to the fact that every glider landing is achieved easily without using any chemical energy; on the other hand, a GA aircraft provides more opportunity to also explore the basic aerobatic manoeuvres — rolls, loops and stall turns. Incidently, 30 minutes thermalling in a glider will also demonstrate the absolute need to coordinate rudder and aileron in every turn — the ailerons on the long slender wings provide a large adverse yaw moment. These intentional spins should not be made from level flight, as described above, but should be made from those flight situations where unintentional spins are most likely to occur; i.e. in climbing and descending turns. These defensive flying lessons will also expose the student to the fact that it is very easy to invoke an accelerated stall during the pull-out after breaking out of the spin, if excessive control column back-pressure is applied — which is an automatic reaction if the ground is rising up to smite you! Spin recovery training in a spin-certified aircraft does invoke a recognition of the behaviour of that aircraft type immediately prior to an incipient spin event, but these warning signs will vary between aircraft types — or even aircraft of the same type. However, knowing how to recover from a stall/spin situation is of no help if it develops at a height that does not provide sufficient height for recovery — circuit height, for example. At such heights the aircraft must always be operated at a safe airspeed (1.5 × Vs1) and restricted to gentle manoeuvres. The 'falling leaf' manoeuvre During World War I, a very simple manoeuvre called the 'falling leaf' was developed as an incipient spin training exercise consisting of repeated entries into incipient spins while the ailerons are held in the neutral position. Although simple, the exercise requires precise timing — so that speed remains at and just above stall — and a good 'feel' on the controls plus very good 'hands and feet' coordination. It involves the initiation of an incipient spin by bringing the aircraft to the point of stall in level flight, then pulling back on the control column whilst applying full rudder. As the wing drops, the control column is moved forward to the neutral position to unstall the wings, then opposite rudder is applied and held, which actions stop the yaw and the incipient spin. Then the control column is pulled back again to stall the wings whilst the rudder is held in the same position; thus an opposite-direction incipient spin is started. Those sequences are repeated so that as the aircraft mushes down, in and out of the stalled condition, it slips from side to side in a series of small arcs, supposedly as a falling leaf may descend. Nowadays the falling leaf is classified as an aerobatic manoeuvre, thus performing the exercise in an ultralight is prohibited. Obviously before any falling leaf exercise is attempted the pilot must receive appropriate spin recovery training. The falling leaf term is also used to describe the technique of 'walking or pedalling down' a stalled aircraft by picking up a dropping wing with opposite rudder and then leaving the rudder applied a little longer than necessary so that the other wing starts to drop. In the latter technique, which is also a good developmental exercise in smooth air, the aircraft shouldn't be allowed to display much lateral movement during the descent. One of Bob Hoover's popular airshow demonstrations, the 'Tennessee Waltz', is a graceful falling leaf manoeuvre. Again, this exercise should not be attempted unless the pilot has appropriate spin recovery training — and ample height because of the substantial height loss in all falling leaf manoeuvres — but all these types of control exercises do provide an excellent means of familiarising yourself with the feel of your aircraft at low speed and its particular stability foibles. Picking up a dropping wing with rudder There is occasionally some debate about the merits of using the secondary effect of rudder to pick up a dropping wing when flying at or near the stall. The dropping wing has not been arrested by roll stability because it is partly or fully stalled. (The reason for proposing use of rudder rather than aileron is because if the dropping wing is near the critical angle of attack the use of aileron will increase the camber of that section of the wing taking it into, or further into, the reducing lift zone of the wings CL curve.) As demonstrated in the falling leaf, using only the rudder to 'pick up' the wing does nothing to remove the stall condition, and excessive input will lead to the opposite wing dropping and the aircraft entering an opposite-direction incipient spin. This technique of picking up a dropping wing with opposite rudder should not be applied during normal stall recovery, unless there is ample height for recovery from an induced spin. The wing must be unstalled by moving the control column forward so that normal aileron control actions can be taken and rudder used to check any yaw. The aircraft manufacturer's recommendations for stall recovery should be followed. But in their absence, the recommended technique in normal stall recovery is always to unstall the wings by easing forward on the control column — which is immediately effective — use sufficient rudder to check any further yaw, at the same time apply full power and then level the wings with aileron. For further information see Standard recovery procedure for all stall types However, when in the final stages of landing, and just above the surface in ground effect (should you want the aircraft to touch down in a stalled condition), gentle application of rudder using opposite yaw to pick up a dropping wing coupled with a slight easing of control column back-pressure may be an alternative to applying power for a go-around. But it depends very much on the particular wing — form, washout, flap setting, slats and slots — on how the stall develops along the wing and on the pilot's knowledge of the particular aircraft. It also depends on how crosswind is being countered. In some aircraft, the use of aileron to pick up the dropped wing will increase induced drag on the lower wing, and the consequent adverse yaw may swing the aircraft towards the ground. The spiral dive As explained above, a spin and a turn are completely different beasts. Also, in a spin, the airspeed is relatively low and constant, the vertical speed is relatively low (2000–4000 fpm) and the angular rotation is fast. In a diving turn or 'spiral dive', the rotation is slower because of the wide (but tightening) turn radius, and the airspeed and height loss both increase rapidly. In the lateral stability section, the possibility of entering a spiral dive condition was mentioned. In a well-developed steep spiral dive — the 'graveyard spiral' — the lift being generated by the wings (and thus the load factor) to provide the centripetal force for the high-speed diving turn, is very high and the turn continues to tighten. The pilot must be very careful in the recovery from a fully established spiral dive, or excessive structural loads will occur. See recovery from a spiral dive. 3.10.7 The stick force gradient An aircraft's control systems must provide the pilot with handling qualities appropriate to the task in hand plus adequate feedback of the aerodynamic forces being generated by the control surfaces — particularly the elevators. To avoid inadvertent airframe overstress it is required that the pilot must always apply an increasing pressure to the control column if increasing the elevator's aerodynamic force and thus the load on the airframe. Increasing back-pressure if the manoeuvre is a turn or a pull-up, increasing forward-pressure if a push-down. That control column pressure requirement is known as the 'stick force gradient' and the pressure applied is specified as the 'stick force per g'. The stick force that can be applied to achieve a particular elevator deflection depends on the length of the control column and the degree of travel available in the fore-and-aft arc; i.e. the stick's mechanical advantage. If the stick travel is short then the force required to deflect the elevators will be greater and the control system will probably feel too sensitive. Also the cg position affects the stick force required to increase the aerodynamic load; an aft cg reduces the stick force, a forward cg position increases the stick force required. To reduce the possibility of inadvertent application of airframe loads exceeding the design positive load limit, the control system must be set up so that the stick force required to reach that limit must be at least a specified minimum value. FAR 23.155 specifies that value as the aircraft's mtow/140 or 15 pounds (6.8 kg), whichever is greater. Take for an example a 600 kg mtow aircraft: 600/140 = 4.3 kg stick force, but as the 6.8 kg minimum is greater, then FAR 23 would require 6.8 kg force as the minimum — for 'control column' systems. However it would not be difficult for the average male to apply a 7 kg one-handed pull on the control column. FAR 23.155 states that the stick force need not be greater than 35 pounds (16 kg). There is a different FAR 23.155 standard for 'control wheel' systems. FAR 23.155 also requires that 'There must be no excessive decrease in the gradient of the curve of stick force versus maneuvering load factor with increasing load factor.' Even design by professionals may not provide a guarantee that the aircraft is safe. Read this United States Federal Aviation Administration special review team report [pdf format] which identified issues with a LSA category aircraft's wing structure, flutter characteristics, stick force gradients, airspeed calibration, and operating limitations. The control system must be set up so that the stick force required to increase load by 1g (the stick force per g) is always greater than a minimum value. Perhaps 1–2 kg force would cover the recreational aircraft range from the lightweight minimum aircraft to the 600 kg light sport aircraft. The lower the value the more sensitive the aircraft is to elevator inputs. If the stick force per g was 2 kg then the pilot would apply a back pressure of 2 kg to increase the load from 1g to 2g for a 60° banked level turn. Once applied that force must be held-on by the pilot; i.e. if the pilot has to ease-off pressure to hold the aircraft in a constant rate 2g level turn then the aircraft is exhibiting signs of instability. Things that are handy to know Top rudder refers to the relative position of the rudder pedals — 'top' being the rudder pedal opposite the lower wing. Thus, if the aircraft is banked and turning to the left, then pressure on the right rudder pedal will apply top (or outside) rudder; pressure on the left rudder pedal will apply bottom (or inside) rudder. Stuff you don't need to know The Wright brothers were the first to realise that control in each of the three axes was necessary for sustained stable flight. They added vertical tail rudders to their canard configuration 1903 Flyer and arranged simultaneous rudder deflection with the wing warping control (the trailing edge of one outer mainplane was pulled down by cords to increase camber while the other was pulled up) thus providing a lateral stability system and countering adverse yaw when initiating a turn. Glenn Curtiss was the first to patent the use of ailerons [but not the first to use] in place of wing warping in the hope of beating the Wright patent for three-axis control/stability systems. A long and bitter patent battle among the United States aircraft manufacturers (that had inhibited U.S. aircraft development and manufacture for 10 years) ended in early 1917 when the U.S. Congress forced all manufacturers into pooling their patents by threatening to seize all their patents. When the USA entered World War 1 in 1917 the nation was forced to purchase 10 940 combat aircraft from Europe while only 5064 mainly training aircraft were supplied to the U.S. army and navy by U.S. manufacturers. View the Wright Brothers 1906 patent. Note that the patent uses the term 'aeroplane' in lieu of 'mainplane' or 'wing'. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.9.1 Concepts of stability and trim The aircraft's response to disturbance is associated with the inherent degree of stability; i.e. self-correction built in by the designer — in each of the three axes — that eventuates without any pilot action. Another condition affecting flight is the aircraft's state of trim — or equilibrium where the net sum of all forces equals zero, i.e. the aerodynamic forces are balanced and the aircraft maintains a steady flight condition when cruising, climbing or descending. Some aircraft can be trimmed by the pilot to fly 'hands off' for straight and level flight, for climb or for descent. But very light aeroplanes generally have to rely on the state of trim built in by the designer and adjusted by the rigger, although most have a rather basic elevator trim device, but no rudder or aileron trim facility. If natural trim is poor — and perhaps it flies with one wing low — inherent stability may maintain equilibrium with that wing-low attitude and not restore the aircraft to a proper wings-level attitude. In which case, the pilot has to maintain a slight but constant control column deflection to hold the wings level, which can be quite annoying. It is desirable that longitudinal trim doesn't change significantly with alterations in power, nor does directional trim change significantly with alterations in airspeed. An aircraft's stability is expressed in relation to each axis: lateral stability — stability in roll, directional stability — stability in yaw and longitudinal stability — stability in pitch. The latter is the most important stability characteristic. Lateral and directional stability have some inter-dependence. Degrees of stability An aircraft will have differing degrees of stability about each axis; here are a few examples: When disturbed a totally stable aircraft will return, more or less immediately, to its trimmed state without pilot intervention; however, such an aircraft is rare — and undesirable. We usually want a sport and recreational aircraft just to be reasonably stable so it is comfortable to fly. If overly stable they tend to be sluggish in manoeuvring and heavy on the controls; i.e. significant control force is required to make it deviate from its trimmed state. If it tends toward instability the pilot has to continually watch the aircraft's attitude and make the restoring inputs, which becomes tiring, particularly when flying by instruments. Some forms of instability make an aircraft unpleasant to fly in a bumpy atmosphere. The normally stable or positively stable aircraft, when disturbed from its trimmed flight state, will — without pilot intervention — commence an initial movement back towards the trimmed flight state but overrun it, then start a series of diminishing damping oscillations about the original flight state. This damping process is usually referred to as dynamic stability (or the tendency over time) and the initial movement back towards the flight state is called static stability. The magnitude of the oscillation and the time taken for the oscillations to completely damp out is another aspect of stability. Unfortunately a statically stable aircraft can be dynamically unstable in that plane; i.e. the oscillations do not damp out. The neutrally dynamically stable aircraft will continue oscillating after disturbance, but the magnitude of those oscillations will neither diminish nor increase. If these were oscillations in pitch, and if there were no other disturbances and the pilot did not intervene, the aircraft would just continue 'porpoising'. The negatively stable or fully unstable aircraft may be statically unstable and never attempt to return towards the trimmed state. Or it can be statically stable but dynamically unstable, where it will continue oscillating after disturbance, with the magnitude of those oscillations getting larger and larger. Significant instability is an undesirable characteristic, except where an extremely manoeuvrable aircraft is needed and the instability can be continually corrected by on-board 'fly-by-wire' computers rather than the pilot — for example, a supersonic air superiority fighter. The best piston-engined WWII day fighters were generally designed to be just stable longitudinally, neutrally stable laterally and positively stable directionally. 3.9.2 Longitudinal stability Longitudinal stability is associated with the restoration of aoa to the trimmed aoa after a disturbance changes it; i.e. if a disturbance pushes the nose up the tailplane will counter with a nose-down pitching moment. In section 6.2 we discussed the provision of a tailplane to act as a horizontal (longitudinal) stabiliser. Before we go any further we need to look at another structural aspect of the airframe. Angle of incidence Angle of incidence is a term that is sometimes mistakenly used as synonymous with wing angle of attack; however, the former cannot be altered in flight except in weight-shift control aircraft (hang gliders and trikes). Angle of incidence, usually just expressed as incidence, is within the province of the aircraft designer who calculates the wing aoa to be employed in the main role for which the aircraft is being designed, probably the aoa in performance cruise mode. The designer might then plan the fuselage-to-wing mounting so that the fuselage is aligned to produce the least drag when the wing is flying at the cruise aoa. Wings that incorporate washout will have differing angles of incidence at the wing root and at the outer section. A notional horizontal datum line is drawn longitudinally through the fuselage, and the angle between that fuselage reference line [FRL] and the wing chord line is the angle of incidence. Incidence should be viewed as the mounting angle of the fuselage rather than the mounting angle of the wings — see 'Stuff you don't need to know'. Incidence may also be called the 'rigger's incidence' or some similar expression carried over from the earlier days of aviation. For ultralight aircraft, incidence is something that should be checked at regular inspections by a qualified person Longitudinal dihedral An angle of incidence is also calculated for the horizontal stabiliser with reference to the FRL. The angular difference between wing and stabiliser angles is called the longitudinal dihedral, although it is probably more correct to say that the longitudinal dihedral is the angular difference between the two surfaces at their zero lift aoa. The angle of the line of thrust is also expressed relative to the FRL. Positive longitudinal dihedral — where the wing incidence is greater than that of the stabiliser — will help control a stall by ensuring that, if the aircraft approaches a stall, the wing will stall before the tail, giving the tail a chance to drop the nose. The tailplane of most very light 3-axis control aeroplanes is mounted in a position where the wing downwash may effect the angle of attack of the tailplane and that downwash angle increases as the wing angle of attack increases. It is the horizontal stabiliser area and moment arm that provides the restoring moment to return aoa to the trimmed state. However, bear in mind that the moment arm, which supplies the restoring leverage and thus the stability, is affected by the cg position. If the cg lies outside its limits, the aircraft will be longitudinally unstable. We learned in section 2.6 that when flying with level wings, at a particular weight, each aoa is associated with a particular IAS. We might as well take advantage of that by arranging the longitudinal dihedral so that the built-in state of trim produces a particular indicated airspeed. In some ultralights a designer/rigger might pick Vbg — best power-off glide speed — as the natural airspeed so that, lacking pilot input, the aircraft will naturally attempt to adjust its aoa to the Vbg aoa, whether power is on or off. Oscillating motions It is possible that an aircraft, properly trimmed for continuing level flight, may develop a 'phugoid' motion if affected by a sharp disturbance. A phugoid cycle is a pitch increase followed by a pitch decrease without any discernible aoa change, i.e. a short climb during which speed decreases and the nose drops into a short descent during which speed increases and the cycle starts again. The aircraft is trading kinetic energy for an increase in the potential energy of height, using the latter to return to the trimmed airspeed in the descent; the cycle time for one oscillation in a very light aircraft might be 20 seconds or so. The oscillating motion issometimes described as 'porpoising'. If the pilot doesn't intervene and the aircraft is phugoid stable the phugoid cycles will damp out after a few diminishing oscillations. If the aircraft is phugoid unstable the oscillations will diverge and the pilot must intervene. The longitudinal dihedral and the tail moment arm affect phugoid stability. 3.9.3 Directional stability Directional stability is associated with the realigning of the longitudinal axis with the flight path (the angle of zero slip) after a disturbance causes the aircraft to yaw out of alignment and produce slip; remember yaw is a rotation about the normal (vertical) axis. In section 6.3 we discussed the provision of a fin to act as a directional stabiliser. The restoring moment — the static stability — provided by the fin is the product of the fin area and the moment arm. The moment arm leverage will vary according to the cg position — the aircraft's balance. The area required for the fin has some dependency on the net sum of all the restoring moments associated with the aircraft fuselage and undercarriage side surfaces fore (negative moments) and aft (positive moments) of the cg. For instance, the Breezy has, except for the pilot's body, very little lateral moment ahead of the cg because of the open frame fuselage; thus a small fin provides all the moment necessary for directional stability. But if the pilot and passenger were enclosed in a cockpit or pod, with a much greater side surface, then the negative moments would be greater and consequently the fin area would have to be greater. If the pilot removes his/her feet from the rudder pedals the rudder, will 'float', aligning itself with the relative airflow and thereby reducing the restoring moment of the fin. The directional stability of very light aircraft with a lot of forward keel area — such as those with a cockpit pod and a 'boom' in place of a rear fuselage — may be 'conditional'; i.e. it is sensitive both to the position of the cg within its normal range and to the amount of sideslip. This is because the negative lateral forces of the pod are very much greater than the positive lateral forces of the boom and fin. Thus, beyond a certain angle of slip the moments change, positive stability is changed to neutral stability and yaw becomes locked in. It might also be associated with the fin stalling at high sideslip angles. The most noticeable symptom to the pilot is aerodynamic rudder overbalance (or 'rudder force reversal' or 'rudder lock') — where the rudder moves to full deflection without any additional pilot input, or doesn't return to the neutral position when the rudder pedal pressure is released, or the pedal force has to be reversed as sideslip angle is increased. It may require significant opposite rudder input, and probably an increase in airspeed, to return to the normal state. The areas of side surface above and below the cg also affect other aspects of stability. The term 'weathercocking' refers to the action of an aircraft, moving on the ground, attempting to swing into wind. It is brought about by the pressure of the wind on the rear keel surfaces, fin and rudder, which cause the aeroplane to pivot about one or both of its main wheels. It is usually more apparent in tailwheel aircraft because of the longer moment arm between the fin and the main wheels; although if a nosewheel aircraft is 'wheelbarrowing' with much of the weight on the nosewheel, then there will be a dangerously long moment arm between the nose wheel pivot point and the fin. 3.9.4 Lateral stability Lateral stability refers to roll stability about the longitudinal axis; in section 4.10 we established that ailerons provide the means whereby the aircraft is rolled in the lateral plane. However, unlike the longitudinal and normal planes where the horizontal and vertical stabilisers provide the restoring moments necessary for pitch and yaw stability, no similar restoring moment device exists in the lateral plane. But let's imagine that some atmospheric disturbance has prompted the aircraft to roll to the left, thus the left wingtip will be moving forward and down, and the right wingtip will be moving forward and up. Now think about the aoa for each wing — the wing that is moving down will be meeting a relative airflow coming from forward and below, and consequently has a greater aoa than the rising wing. A greater aoa, with the same airspeed, means more lift generated on the downgoing side and thus the left wing will stop going further down or perhaps even rise a little, although pilot action is usually needed to get back to a wings level state. This damping of the roll is known as lateral damping. So roll stability, except when at or very close to the stall, is intrinsic to practically all single-engined light aircraft. (When the aircraft is flying close to the stall, the aoa of the downgoing wing could exceed the critical aoa and thus stall, which will exacerbate the wing drop and might lead to an incipient spin condition. See the stall/spin phenomenon.) But — and there always seems to be a 'but' — when the aircraft is banked, other forces come into play and affect the process. If you re-examine the turn forces diagram in the manoeuvring forces module, you will see that when an aircraft is banked the lift vector has a substantial sideways component; in fact, for bank angles above 45°, that sideways force is greater than weight. So we can say that any time the aircraft is banked, with the rudder and elevators in the neutral position, an additional force will initiate a movement in the direction of bank; i.e. creating a slip. We know from the section 7.3 that the aircraft's directional stability will then yaw the nose to negate the slip and the yaw initiates a turn, which will continue as long as the same bank angle is maintained. There are several design features that stop the slip and level the wings, thus promoting lateral stability. For instance, placing the wing as high as possible above the cg increases so-called 'pendular stability', (The stability due to the high wing is not really pendulum stability such as that applicable to powered parachutes.) Wing dihedral* is usually employed with low-wing monoplanes (and to a lesser degree of tilt with high wings), where the wings are tilted up from the wing root a few degrees. A swept-back wing format is used with trikes. Another design method is anhedral, where the wings are angled down from the wing root, but it is unlikely to be used in light aircraft, although the powered parachute wing utilises an anhedral arc for stability. (*'Dihedral' is a mathematical term denoting the angle between two intersecting planes.) Spiral instability An aircraft with positive spiral stability tends to roll out of a turn by itself if the controls are centred. Some light aircraft with little or no wing dihedral and a large fin tend to have strong static directional stability but are not so stable laterally. If slip is introduced by turbulence or by the pilot, such aircraft — left to their own devices — will gradually start to bank and turn — with increasing slip and nose drop — and hence increasing turn rate and rapid increase in height loss. Neutral spiral stability is the usual aim of the designer. The turning process starts slowly in aircraft with slight spiral instability but leads to spiral divergence which, if allowed to continue and given sufficient height, will accelerate into a high-speed spiral dive. This often occurs when a pilot without an instrument flight rating strays into cloud where all visual cues are lost. In that condition it is known as the 'graveyard spiral'. Inadvertent entry into a fatal spiral dive, leading to inflight breakup, can happen even with experienced IFR pilots, see this Australian Transport Safety Bureau report. It is evident that directional stability and lateral stability are coupled (i.e. rotation about one axis prompts rotation about the other) and to produce a balanced turn; i.e. with no slip or skid, the aileron, rudder and elevator control movements and pressures must be balanced and coordinated. Dutch roll Induced motion in the lateral plane generally brings about a coupled motion in the directional plane, and vice versa. Dutch roll is a phenomenon in level flight where a disturbance causes a combined yaw and roll followed by a return to the level flight condition then a yaw and roll to the other side: the oscillations continuing until damped out. In a very light aircraft the time for each cycle might be 5 to 10 seconds. The motion is quite uncomfortable, viewed from the cockpit the wingtips complete a circular motion against the horizon as does the nose. Pilot intervention is by use of rudder. 3.9.5 Trim and thrust We have covered above the reaction of the aircraft to changes in relative airflow whether induced by the pilot or minor atmospheric turbulence. We know from sections 1.8 and 1.9 that if an aircraft is properly trimmed for cruise flight and we increase thrust then it will climb; and if we reduce thrust it will descend. But how this eventuates is not at all straightforward. The reaction to changing power, without the pilot touching the control column, depends on whether the cg is above, below or inline with the line of thrust; in the Breezy, the cg is below the thrust line. The thrust line is best located so that it passes close to the vertical cg position to minimise the initial pitching moments associated with power changes. The placement of the horizontal and vertical stabilisers, in relation to the propeller slipstream and to the wing downwash, affects flight performance and particularly flight at slow speeds — because then the total air velocity within the slipstream tube is nearly double that outside the tube; also the slipstream is rotating, and will thus impart a sideways moment to the fuselage and vertical stabiliser. Effects on individual aircraft types vary according to the designer's inbuilt compensations: for example, if the horizontal stabiliser operates in the wing root downwash airflow, then when the wing root stalls and the downwash becomes turbulent the stabiliser might undergo an abrupt change in aoa (and thus in its stability restoring moment). Or if the horizontal stabiliser operates substantially outside the downwash but if it is in the path of the turbulent flow from the stalled wing, it will then lose part of its aerodynamic force. If a modification is made to that design, even a seemingly minor change, the consequential effect on stability may be quite surprising. To illustrate the point, I suggest you read an "airworthiness report regarding (among other factors contributing to general stability problems) a small change made in relocating the exhaust manifold of a Thruster that, at a particular aoa, promoted turbulent flow over the upper wing surface, which then extended to the horizontal stabiliser, and reduced the stabilising moment imparted by that surface. Stuff you don't need to know The term 'decalage' (French = gap or shift forward/back) relates to the difference in the angles of incidence of the upper and lower mainplanes of a biplane. Decalage is now occasionally used as synonymous with longitudinal dihedral. The angle of incidence has some effect on the pilot's view over the nose. A very few naval aircraft designs have included 'variable incidence wings' where the angle of incidence could be changed by the pilot during flight, within a range of say 2–15°, using electric motors. Such aircraft included leading edge slats as a high-lift device.The idea was to take full advantage of the high maximum CL and consequent low speed, during the landing approach, without having the fuselage cocked up at a high angle blocking the view. As aoa increased and the aircraft slowed, the pilot wound the fuselage down, so that it remained more or less level during the approach and thus provided a better view of the flight deck! Variable incidence wings were also used with one of the post-WWII Supermarine amphibian designs. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.8.1 Airframe basics The empennage A monoplane has a single wing, or a left and right pair of wings — the port and starboard mainplanes. A biplane has two sets of wings mounted one above the other. The engine and propeller may be in front of the wing (a 'tractor' configuration) or behind the wing (a 'pusher' configuration). It must have some sort of pilot/passenger seating (usually enclosed in a cockpit or pod, which will have either tandem or side-by-side seating and is referred to as the 'occupant zone'); fuel tank(s); and a rigid structure mounting the engine/propeller, the wheels or undercarriage, and the cockpit, so that their weight is supported by the wing main spar(s) and the propeller thrust reaction is transmitted to the body and wings. The photograph of Mick McCann's 'Breezy' shows a basic high-wing monoplane, pusher-engine configuration with tandem pilot/passenger seating, nose-wheel undercarriage and an open-frame, welded tubular-steel fuselage — the aft part of which is upswept, so that the aircraft's attitude in pitch can be adjusted during take-off and landing without the tail striking the ground. Also the arrangement keeps the rear stabilising and control surfaces within the energetic airflow of the propeller slipstream. The term fuselage is derived from an old French word meaning a tapered 'spindle' used for manually weaving yarn. The 'Breezy' has no refinements for comfort — or for drag reduction. The fuel tank is not discernible in the photograph but is small and close to the engine. Attached to the rear fuselage are the horizontal stabiliser and elevators, plus the vertical stabiliser or fin and the rudder, together forming the pitch and yaw stabilising and control mechanisms — the tail assembly or empennage. The latter term is derived from a French word meaning to feather an arrow; maybe that is why some people refer to the empennage as the 'tailfeathers'. The horizontal stabiliser and elevators are referred to as the 'tailplane'. Moments and couples The moment of a force or the torque is a measure of the rotational effect produced by a force acting about — or with respect to — a fulcrum, axis, centre of mass (cg) or aerodynamic centre. Its magnitude is the product (in newton metres) of the force (N) and the length (m) of the arm (the leverage) from the pivotal point to the line of action of the force. The moment will act in a particular direction, for example, as we saw in the 'Aerofoils and wings' module, the pitching moment of a cambered wing produces a nose-down torque. The forces generated by the tailplane control surfaces are dependent on the stabiliser area, the control surface area, the length of the tail arm to the cg, the control surface deflection and the airspeed. Only deflection and airspeed are controlled by the pilot. Two equal and opposite forces acting parallel to each other, but separated, form a couple. The rotational effect or moment of a couple is the product of one force and the perpendicular distance between them. The ailerons, for example, form a couple when deflected. 3.8.2 Tailplane Horizontal stabiliser The existence of the wing pitching moment makes the wing inherently unstable. To overcome this problem, it is necessary to couple it with another aerodynamic moment about the lateral or pitch axis — opposing the wing pitching moment — that will balance that moment at an airspeed selected by the designer. The moment of a force is the arm length multiplied by the force; so the longer the tail arm, the smaller the aerodynamic force required. The standard solution is to extend the fuselage rearwards so that a horizontal stabiliser can be mounted at a distance from the cg; note the Breezy's very long tail arm – between the cg and the small horizontal stabiliser. The horizontal stabiliser is usually a lift-generating surface — or 'plane' — mounted so that the aerodynamic force it generates acts in the opposite direction to the lift from the mainplane, i.e. generally downwards. The plane could incorporate a cambered aerofoil with the cambered surface underneath, or perhaps a symmetrical aerofoil, or even just a flat plate — as the Breezy's appears to be. The symmetrical aerofoil and the flat plate would both be mounted at a negative incidence to produce the downward force. The end result is that the net pitching moment of the mainplane and tailplane couple is zero at a particular geometric aoa of the main wing; that aoa would equate with a speed selected by the designer — usually the designed cruise speed or perhaps the engine-off glide speed. The fuselage may also produce pitching moments that must be balanced by the stabiliser. As the horizontal stabiliser is usually designed to produce negative lift, then the wing must fly at a slightly greater aoa to provide additional lift, so that the net aircraft lift balances weight. Elevators The pilot must be able to initiate and hold aoa changes for airspeed adjustments, manoeuvres (accelerations) in the pitching plane (pull-ups, turns, push-downs) and adjustments of aircraft attitude relative to the airfield or alighting area surface during take-off and landing. The elevators — hinged to the trailing edge of the horizontal stabiliser so that they may be deflected up or down — are the control surfaces that enable controlled changes in wing aoa. Elevators are aerodynamically similar to the ailerons, but move in unison rather than differentially. The elevators are linked, via control rods or cables, to forward/backward movement of the control column, so the pilot can, in effect, increase or decrease the camber of the stabiliser–elevator combination. Camber changes will alter the magnitude and direction of the aerodynamic reactions generated by the stabiliser–elevator, and the changed forces impart a pitching moment in the longitudinal plane. This pitching moment rotates the aircraft about its lateral axis, initiating the change in wing aoa. Once the new aoa is established, the pitch moment returns to zero and the aircraft will hold that aoa — provided the elevators are held in the deflected position by the pilot or a trim device — thereby controlling airspeed for a given power setting. Backward movement of the control column raises the elevators and the aircraft's nose pitches up; forward movement lowers the elevators and the aircraft's nose pitches down. The force able to be exerted via the elevators is the most significant control force. The 'up' and 'down' terms in pitch are not relative to the horizon but to the original flight path in the aircraft's longitudinal plane. A stabilator is an 'all-moving' or 'all-flying' tailplane combining the horizontal stabiliser and elevator providing similar force with a lesser deflection, thus less drag. Sometimes used in higher speed light aircraft but rarely in very light aircraft. There may be some net advantages in mounting the stabiliser and elevators in front of the wing — a canard — but such arrangements are rather rare amongst very light aircraft. 3.8.3 Vertical stabiliser and rudder Because of drag and other effects, aircraft perform much better if their longitudinal axis is accurately aligned, in plan view, with the flight path. If unaligned, the aircraft velocity will have both a forward component and a slight lateral component, and the relative airflow — the flight path — will not be aligned with the longitudinal axis. Such bodily sideways (translational) movement along the lateral axis is called slip or sideslip or skid. The skid term is generally associated with excess 'bottom' rudder and skidding out in a turn, as a road vehicle might. Thus, some means is required to ensure that if the horizontal direction of the relative airflow is changed (i.e. the aircraft acquires slip because of a minor disturbance) then the aircraft will automatically yaw — rotate itself about its normal axis — to realign its longitudinal axis with the airflow, so that the sum of all the lateral moments — fore and aft of the cg — equals zero. The long-established means is to use a fin, or vertical stabiliser, mounted at the rear of the aircraft, that has an aerofoil section — usually symmetrical — or is just a flat plate. The fin applies the restoring moment to realign the longitudinal axis with the airflow. That moment does not realign the aircraft with its original flight path; after restoring alignment with the relative airflow, the aircraft may be aligned with a different flight path, depending on the amount of original displacement. The fin is often angled away from the aircraft's longitudinal axis by a few degrees. This offset creates an aerodynamic force that compensates for the rotating propeller slipstream applying a force to one side of the fin. The rudder is the control surface hinged to the fin and is the lateral plane equivalent of the elevators; though the rudder is operated by the pilot's rudder pedals rather than the control column. Pressure on the left pedal causes the rudder to deflect to the left, so that the fin/rudder act together as a cambered aerofoil to produce an aerodynamic force that pushes the tail to the right — and consequently the nose swings left; i.e. the aircraft yaws left. (Yaw is an old nautical term associated with the motion of the sea swinging the bow off-course.) The amount of yaw, at a given airspeed, is dependent on the degree of rudder deflection. (But, of course, it is primarily dependent on the tail moment arm and rudder area.) The aircraft will continue yawing if the rudder deflection is held by the pilot, but as the aircraft turns (i.e. it is rotating about its normal or vertical axis while moving forward), the wing on the outside of the turn must be moving slightly faster than the inner wing and thus generates more lift. The increased lift will raise the outer wing and the aircraft will enter a banked turn, but will tend to skid out because the bank angle will not be correct for the turn. Only one bank angle will produce the desired radius or rate of turn for a particular airspeed. Note the Breezy's small fin with its relatively large rudder. The pilot's feet are on the pedals linked to the rudder and he is holding the control column — linked to the ailerons and the elevators — with one hand. The other hand is probably holding the engine throttle lever. The rudder initiates yaw about the normal axis; the ailerons initiate roll about the longitudinal axis; the elevators initiate pitch movement about the lateral axis. 3.8.4 Control balance Aerodynamic balance Aircraft designers try to impart a good 'feel' to the controls so that the pilot finds they are not too 'heavy' or too 'light' to operate through most of the speed range. So, the elevators and rudder are usually fitted with some sort of aerodynamic balance, which puts part of the control surface forward of the hinge line. Such devices might be inset hinge balances, leading-edge balances or control horns that reduce the hinge moments needed to deflect the control surface. Mass balance Control surfaces need to be hinged near the leading edge, the centre of their mass will be well aft of the hinge line; i.e. the mass of the control is not statically balanced. That, combined with the necessary elasticity of aircraft structures, leads to a control flutter problem. This might occur with mass unbalanced control surfaces at any speed, but particularly with ailerons at high speed. Flutter has the potential to lead to structural failure. The prime solution to the mechanical unbalance and the flutter problems is for the manufacturer to accurately balance the mass of the control surface by inserting weights forward of the hinge line usually within the hinge insets or the control horn. This — known as mass balance — increases the stability of the control surface and ensures that accelerations don't deflect the control surface. Notes for scratch-builders You may occasionally come across the terms tail volume and tail volume ratio. The horizontal tail volume is the surface area of the horizontal stabiliser plus elevators multiplied by the length of the moment arm of the horizontal stabiliser measured from the wing MAC quarter chord to the horizontal tail MAC quarter chord. The horizontal tail volume ratio or tail volume coefficient is the tail volume divided by the product of wing area and wing MAC. Tail volume ratio is usually in the range 0.35 to 0.45 for minimum aircraft and 0.45 to 0.55 for aircraft of composite construction — when the units of measurement are feet. The higher the coefficient value, the more stable the aircraft. There is a similar equation for the vertical stabiliser and rudder, but the divisor is the product of wing area and wing span. Such ratios are of interest to an aircraft designer, as there is a linear relationship between tail moment or tail area, and stability — doubling the tail moment or the tail area doubles the static stability and quadruples the dynamic stability. The aspect ratio of the tail affects tailplane pitching moments; for a given area higher aspect ratio produces less induced drag and hence the lift component of the aerodynamic force is higher. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.7.1 Engine power output Engine power equals the product of force and speed. Torque is the rotational force acting about the engine crankshaft multiplied by the moment arm; i.e. it is the product of the firing stroke in the cylinder and the radius of the crank to which the connecting rod is attached. The bigger the cylinder the bigger the rotational force — the 'bang'. Engine speed is measured in crankshaft revolutions per minute [rpm]. In the 'Manoeuvring forces' module we discussed the power required for various flight conditions, and looked at power required/power available curves and the effect of altitude on power output. It may be appropriate to review section 1.7 of that module. Normally aspirated aero engines The maximum power that can be developed, in the cylinders of a particular piston engine, increases or decreases directly with the density of the air in the intake manifold, and air density decreases as altitude increases — or temperature increases. See the atmospheric density and the International Standard Atmosphere sections in the 'Airspeed and the properties of air' module. Thus, the full throttle power output of a normally aspirated engine — one that relies solely on the ambient atmospheric density — decreases as operating altitude increases. The diagram in section 1.7 shows how maximum brake horse-power [bhp], delivered at full throttle in a normally aspirated engine, decreases with altitude. A 100 hp engine operating at 65% power will be delivering 65 hp. Power produced is proportional to the air density at the intake manifold, the cylinder displacement and compression ratio, the number of cylinders, and the rpm. Of those items, only the air density at the intake manifold and the engine rpm alter, or can be altered, during flight. (With a normally aspirated engine and a propeller whose pitch is not variable in flight, the throttle controls manifold pressure, which then determines rpm.) A traditional four-stroke light aircraft engine, such as the Lycoming O-235, has an individual cylinder displacement of 950 cc, a compression ratio of 7:1 and a maximum design speed of 2600 rpm, at which its rated 110 bhp is produced — in sea-level ISA conditions. The Rotax 912, the most common lightweight four-cylinder aero-engine, utilises an individual cylinder displacement of only 300 cc, a compression ratio of 9:1, but doubles the maximum design speed to 5500 rpm to achieve its rated 100 bhp. The lightweight Jabiru 2200 utilises an individual cylinder displacement of 550 cc, a compression ratio around 8:1 and a maximum design speed of 3300 rpm to achieve its rated 80 hp. The three engines mentioned are all horizontally opposed, four-stroke and four-cylinder; a popular configuration providing a fully balanced engine that doesn't require crankshaft balance weights. Engines are often described in terms of 'total capacity' (cylinder displacement by number of cylinders) in litres or cubic centimetres. Thus, the Lycoming O-235 is 3.8 litres or 3800 cc (235 cubic inches), the Rotax 912 is 1.2 litres and the Jabiru 2200 is 2.2 litres. Most engines used in ultralights tend to be around 30% lighter (in terms of weight per rated hp) than the ubiquitous Lycoming and Continental piston engines used in general aviation aircraft. Thus, they are cheaper to manufacture but less robust, with a consequent shorter time between overhaul [TBO]. Although aero-engines can quite happily operate continually at their rated power, doing so is not good practice. It is uneconomical in terms of fuel efficiency, but — more importantly — it may shorten engine life, if engine operating temperatures and pressures are exceeded. Normally the maximum — and optimum — power setting for continuous cruise operation is 75% of rated power. Turbo-charged aero-engines The volumetric efficiency (i.e. the cylinder-filling capability) of an engine can be improved by increasing the density of the fuel/air charge delivered to the cylinders by compressing the air in the atmospheric intake manifold. This process is supercharging and develops more torque at all engine speeds. The compressor is usually a lightweight centrifugal impeller driven by a gas turbine that utilises the otherwise wasted energy of the engine exhaust gases. Such a system is a turbine-powered supercharger, usually described as a turbocharger. An oil pressure-driven butterfly valve or waste gate is incorporated within the exhaust manifold system, automatically adjusting — according to the pressure within the intake manifold — to allow all, or a portion, of the exhaust gases to bypass the turbine; thus continually maintaining the system within the designed operating limits. There is a slight penalty in that turbocharging also increases the temperature of the charge. This consequently decreases the achievable density and possibly leads to detonation, unless a charge cooling device — an intercooler — is incorporated between the compressor and the cylinders. For some information on mechanically powered supercharging, read this magazine article. Turbocharging may be used to increase the sea-level rated power of the engine, but the use of that full throttle power at low altitudes would normally be limited to short periods because of engine temperature limitations. The big advantage is the increase in power available at altitude. The diagram plots the power achieved (percentage of rated power) at full throttle, in ISA standard conditions, for a normally aspirated engine and the turbocharged version. The turbocharged engine can maintain its rated power from sea-level up to the 'critical altitude', probably around 6000 or 7000 feet, after which it will decrease. The waste gate would probably be fully open at sea-level and then start closing as altitude increases — so that it would be fully closed at, and above, the critical altitude. Turbocharging raises the service ceiling of the aircraft. The service ceiling is the ISA altitude at which the aircraft's best rate of climb (from an extended climb starting at MTOW and unassisted by any atmospheric phenomena) drops below 100 feet per minute — regarded as the minimum useful climb rate. This should be the aircraft's ceiling quoted by the manufacturer. The Rotax 914 series 115 hp turbocharged engines are often regarded as just being suitable for ultralight aircraft. However, those engines power the Predator RQ1/MQ1, unmanned aerial reconnaissance and surveillance vehicles, used so successfully in the Afghanistan and Iraq campaigns of recent years. The Predators have a maximum take-off weight around 1000 kg, cruise around 90 knots, normal mission duration around 20 hours — but could operate for 40 hours — and service ceiling of 25 000 feet. They often carried two 50 kg Hellfire missiles for attacking acquired targets — they also need 5000 feet of paved runway for take-off. Two-stroke aero-engines The lower power (say, up to 65 hp) engines used in ultralight aircraft are usually two-stroke engines, although the half-VW four-stroke auto engine conversions are around 40 hp. Two-strokes don't have very good volumetric efficiency, and the engine is generally efficient only in the upper 30% of its rpm and throttle opening range. In fact, ultralight two-strokes tend to run very roughly at speeds below 2500 rpm and achieve their rated power at rotational speeds in the 5500 to 6500 rpm range. The three most common two-strokes are two-cylinder models with individual cylinder displacements around 250 cc; they achieve their rated power at 6500 rpm and 75% power at around 5300 rpm. Fuel efficiency drops off very quickly as rpm is reduced below the 75 % power figure; see the Rotax two-stroke engine operator's manual section 10.2. Gearing or belt reduction is used to improve the torque delivered to the propeller shaft while also reducing the rpm to something more suitable for the propeller. The torque increases because of the larger rotational radius of the driven gear. The big advantage with two-stroke engines is their mechanical simplicity, and consequent weight and cost saving, because they lack the camshaft and associated valve train of the four-strokes. Some very small (15 hp) two-strokes are used to power self-launching powered hang-gliders. Between 1999 and 2003, there were 98 engine failures reported to RA-Aus; 39 were two-stroke engines and 59 were four-stroke. It is estimated at that time about 65% of the ultralight fleet, of some 1800–2000 aircraft, were equipped with two-strokes. It would appear during that period the two-strokes were more reliable than the lightweight four-stroke aero-engines, though the development of lightweight four-strokes was then not as far along the learning curve as two-stroke development. 3.7.2 Propeller power output Propeller efficiency An aircraft engine supplies energy, in the form of rotational power, to the propeller shaft. The propeller converts the rotational power to thrust power, either pulling the aircraft along behind it (a tractor installation) or pushing the aircraft in front of it (a pusher installation). The pusher installations are usually the only options when the engine is mounted on a carriage or cart rather than a fuselage structure. That option is the standard for trikes, nanolights and gyroplanes. The 10 hp or so engines attached to the 'backpack' harnesses of powered hang gliders and powered paragliders are, of course, pushers. The major problem with pusher propellers in very light aircraft is the avoidance of something from the occupant zone — just in front of the engine — moving through the propeller disc or being entangled in it. The propeller accelerates a tube of air, with much the same diameter as the propeller disc; i.e. it adds momentum to the tube of air and the reaction force propels the aircraft forward. The velocity of this accelerated airstream (the slipstream) has both rotational and rearward components. Momentum = mass × velocity, so if the mass of air passing through per second is increased by increasing the diameter of the propeller, the rearward velocity imparted can be decreased but still produce the same rearward or axial momentum. The rate at which axial momentum is imparted to the air equates with thrust. Propeller efficiency is the ratio of the thrust power (thrust × aircraft forward speed) output to the engine power input. The work done (the energy expended) by the propeller is the kinetic energy imparted to the slipstream = ½mv² joules (if mass is in kilograms and v in metres per second), so less energy is expended if the mass is increased and the velocity decreased. Using a simplified static thrust example, if m = 10 kg and v = 100 m/s, then the momentum is 1000 kg·m/s and energy expended is ½ × 10 × 100² = 50 kJ. But if the values for m and v are interchanged (i.e. m = 100 kg and v = 10 m/s) the momentum will still be the same but the energy expended will be decreased substantially; i.e. ½ × 100 × 10² = 5 kJ. Thus, the most efficient system is to utilise the greatest propeller diameter possible — limited by: the stress effects on the engine (the gyroscopic moments increase exponentially with diameter; see below) ground clearance requirements in worst conditions (e.g. heavy landing and deflated tyre) propeller blade strength – centrifugal forces are extremely high, much greater than aerodynamic forces, even a lightweight blade would be experiencing forces around 2500g. blade tip speed. When a propeller is rotating, the speed at any point on a blade is the product of the rpm and the distance of that point from the hub, and thus the speed at the propeller tip is the greatest. Compressibility constraints dictate that the speed at the blade tips should not exceed about Mach 0.85 — 560 knots or 290 m/s at sea-level. But significant compressibility effects become evident at 250 m/s and, if the propeller is close to the pilot, the noise may be extremely uncomfortable. So, for comfort, tip speed is usually in the range 200–240 m/s. it is not only the aircraft occupants who must be comfortable with the noise, there are restrictions on engine and propeller noise in the vicinity of aerodromes – see the aircraft noise regulations. For light aircraft engine/propeller systems, it is usual to restrict propeller speed to less than 3500 rpm; so, the high rpm engines must incorporate a gear-driven or belt-driven propeller speed reduction unit [PSRU] between the crankshaft and the propeller shaft. The rotational speed of the fixed-pitch propeller depends on the pitch of the blades, the power supplied to the propeller and the aircraft velocity. Propeller blade area is an important consideration in propeller design and choice. Blade aspect ratio is usually maintained around 6–8; so, with a limited propeller diameter, blade area can only be increased by increasing the number of blades. Matching engine and propeller Propellers must be carefully matched with the characteristics of the airframe, engine and reduction gear to which they are mated. The engine must be neither underloaded nor overloaded. At best, a mismatch could make the engine and aircraft incapable of delivering its designed performance, or create the situation where the engine cannot be opened up to full throttle because the lack of load (see the following paragraph) would take the rpm beyond the red-line limit, or it could result in crankshaft or crankcase fracture. At worst, a mismatch could lead to torsional vibration or propeller blade destruction induced by centrifugal force. This can readily cause the engine to dismount from the airframe and lead to consequent total loss of the aircraft. When discussing the power required curve it was noted that power required is proportional to aircraft velocity cubed. Similarly, the power delivered by a propeller varies in accordance with rpm cubed (if everything else is kept constant). Thus, the load on the propeller may be substantially increased just with a relatively minor further increase in rpm when operating at high rpm, which can lead to loss of the blades. Note that centrifugal forces on the blades change in accordance with the rpm squared. Note: The load on the engine is the propeller torque. When the aircraft is stationary, with the engine throttle wide open, the propeller torque and the static thrust generated (i.e. the efficiency of the engine and the propeller combination) depend on the propeller pitch. If the pitch is zero or slightly negative, the static thrust will be zero and the propeller torque will be very low so that the engine will race — overspeed — and lose power because of inefficient cylinder charging, etc. On the other hand, if the pilot is able to set the prop to a more negative pitch, then reverse thrust will be generated together with sufficient torque to maintain constant engine rpm and the aircraft will move backward. If the pitch is 'fine' (low aoa), the propeller will generate near maximum static thrust and sufficient torque to maintain high engine rpm, thus delivering ample power to the propeller shaft. This is the ideal situation to get the aircraft rolling for take-off and climb-out. If the pitch is very 'coarse' (high aoa), then static thrust is low but propeller torque is very high, which will slow the engine. This is the worst situation for take-off — the aircraft will move forward sluggishly and, hopefully, never reach take-off speed. For an interesting article on ground testing of aircraft engines for power output, read "Testing one, two three" in the July-August 2002 issue of 'Flight Safety Australia' magazine. When an aircraft with a fixed-pitch propeller is flying the back of the power curve (i.e. an increasing thrust power output is needed as the airspeed decreases), the propeller efficiency will decrease as airspeed decreases, while the increasing propeller torque will be slowing the engine power. Thus, it may be difficult to arrest any sink that develops at low speeds — as might be experienced on the approach to a short-field landing. However, even with an apparently well-matched engine/propeller combination, there may be a certain rpm range (or ranges) where the frequency of a particular engine vibration resonates, with some natural frequency of the propeller, to produce an intrusive vibration and a potentially damaging stress cycle. In such aircraft, that rpm range or ranges is (or should be) indicated as a yellow, perhaps red, arc on the face of the engine tachometer. Rpm settings within those ranges should not be used. Any gyroscopic moment induced depends on the rate of change in aircraft pitch or yaw, and the rotational speed and moment of inertia of the propeller. Its mass moment of inertia depends on propeller mass and diameter. The gyroscopic loads are transferred to the airframe via the engine crankshaft, crankcase and mountings. Under some conditions, gyroscopic loads may lead to crankshaft/crankcase failures. See 'The Fox story'. The failure conditions usually identified are the use of a propeller of excessive diameter (the moment of inertia increases exponentially with diameter) possibly combined with an excessive 'overhung' moment — the distance from the propeller cg to the engine. Excessive gyroscopic loads may also be placed on the crankshaft/crankcase by using brake, rudder and a burst of throttle to swing an aircraft rapidly when taxiing. The flight conditions that follow propeller blade failure cannot be simulated in training, but an extreme out-of-balance condition (loss of one blade for example) can very quickly shake the engine from its mountings. 3.7.3 Propeller types The following is a copy of a document authored by Marcus Graney and published on the web site of the New Zealand manufacturer of Airmaster propellers. I have added the notes presented in italic. ... JB The most common type of propeller in sport aviation is the fixed-pitch propeller. Although cheap, this is one of the crudest propulsion devices you could use, and has been superseded by a variety of more advanced options, now readily available on the market. But, how do you know how each type of propeller operates and what advantages the different types offer? How are you going to choose between the different types available for your aircraft, especially considering that a more capable propeller is also more expensive? There are four common families of propeller, which I will introduce to you. They are fixed-pitch, ground-adjustable, inflight-adjustable and constant-speed. The last two are both examples of variable-pitch propellers. In order to appreciate the advantages which are characteristic of the different families of propeller, we must first consider the most fundamental characteristic of a propeller — the pitch. Pitch is important, as it is the manner in which pitch is controlled that allows us to differentiate between one family of propeller and another. A useful analogy when considering the affect of pitch is that of an automobile gearbox. By comparing a propeller's pitch to a gear ratio, and considering the function of a gearbox, we will gain an appreciation of the different families of propellers. What is pitch? Propeller theory includes a variety of concepts that may at times be called pitch. Pitch can refer to the blade angle with respect to a flat plane, the distance that a propeller will advance through the air for each rotation or the amount of "bite" that the blade has on the air. Essentially these concepts all describe the same thing. To use our automobile analogy, pitch is like the gear ratio of the gearbox. The important thing to note with pitch, is that it is available in a wide variety of degrees, or 'amounts', much like different gear ratios. To demonstrate, consider the following examples: A fine pitch propeller has a low blade angle, will try to move forward a small distance through the air with each rotation, and will take a 'small' bite of the air. It requires relatively low power to rotate, allowing high propeller speed to be developed, but achieving only limited airspeed. This is like having a low gear in your automobile. A coarse pitch propeller has a high blade angle, will try to advance a long distance through the air with each rotation, and will take a big 'bite' of the air. It requires greater power to rotate, limiting the propeller speed that can be developed, but achieving high airspeeds. This is like having a high gear in your automobile. Pitch and the different families of propellers As we saw above, pitch is a key element in the description of propellers (along with other factors such as diameter and blade area). When considering the four families of propellers it is useful to start with the simple fixed-pitch propeller, and look at the enhancements in pitch control that are gained as we progress through each family to the most advanced, the constant-speed propeller. Fixed-pitch propeller With a fixed-pitch propeller, the pitch of the propeller is fixed from manufacture. The performance of your aircraft is determined on the day your propeller is fitted, and is going to be limited within the constraints of the propeller. An analogy with an automobile is as though you had only one gear. Often when choosing a fixed-pitch propeller for your aircraft, manufacturers give you a choice of either a climb or a cruise prop. A climb propeller has a relatively fine pitch and a cruise propeller has a relatively coarse pitch. This is like a car manufacturer giving you a choice of a low or a high gear. Either you will be really slow off the mark, or your engine is going to have to be red-lined to get anywhere at a reasonable speed. Ground-adjustable propeller Many propellers manufactured and sold for ultralight and experimental aircraft are ground-adjustable. These propellers have the advantage of being able to have their pitch set before each flight if required, taking into account the type of flying you intend to do. More usually however they are used as a low cost way to try out various pitches and settle on the propeller pitch that best suits your aircraft and your style of flying. This can be compared to having a gearbox in your car that you can only change before you set out on your journey. Variable-pitch propeller With a variable-pitch propeller, you really have choices. To use the automobile analogy again, your car now has a real gearbox that you can change gear with on the go. (I hope that your car can do this at least!) In addition, rather than being limited to 4 or 5 gears, you can utilise any pitch along the continuum from maximum to minimum. The pitch of the propeller may be controlled in flight to provide improved performance in each phase of flight. Typically you would take-off in a fine pitch (low gear) allowing your engine to develop reasonable revs, before increasing the pitch (change up gears) as you accelerated to your cruising speed. You'll end up with the propeller at a relatively coarse pitch, (high gear) allowing the miles to pass beneath you at a rapid rate, while your engine is gently ticking over at a comfortable speed. This feature of a variable-pitch propeller will provide you with performance advantages, including: Reduced take-off roll and improved climb performance. Fine pitch allows the engine to reach maximum speed and hence maximum power at low airspeeds. Vital for take-off, climb, and for a go-around on landing. Improved fuel efficiency and greater range. Coarse pitch allows the desired aircraft speed to be maintained with a lower throttle setting and slower propeller speed, so maintaining efficiency and improving range. Higher top speed. Coarse pitch will ensure your engine does not overspeed while the propeller absorbs high power, producing a higher top speed. Steeper descent and shorter landing roll. With a fine pitch and low throttle setting, a slow turning propeller is able to add to the aircraft's drag, so slowing the aircraft quicker on landing. Variable-pitch propellers actually come in a variety of versions. These different versions refer to the different ways that they are controlled, and include: Two-position propeller. Inflight-adjustable propeller. Automatic propeller. Constant-speed propeller. A couple of these are now of historic interest only, so lets concentrate on the two most common options these days; the inflight-adjustable operation and the constant-speed propeller. The inflight-adjustable propeller allows the pilot to directly vary the pitch of the propeller to the desired setting. Combined with the throttle control, this control allows a wide variety of power settings to be achieved. A range of airspeeds can be maintained while keeping the engine speed within limits. While rare in larger aircraft, the inflight-adjustable propeller is the most common type of variable-pitch propeller that is encountered in sport aviation. When operated in manual mode, the Airmaster propeller is an example of an inflight-adjustable propeller. Constant-speed propeller The constant-speed propeller is a special case of variable pitch, which is considered in a family of its own, and offers particular operating benefits. With constant speed control, the pitch of the variable-pitch propeller is changed automatically by a governor. After the pilot sets the desired engine/propeller speed with the propeller speed control, the governor acts to keep the propeller speed at the same value. If the governor detects the propeller speed increasing, it increases the pitch a little to bring the speed back within limits. If the governor detects the propeller speed decreasing, it decreases the pitch a little to bring the speed again back within limits. This operation may be compared to an automatic gearbox in an automobile, where the gears are changed automatically to keep the engine operating at a reasonable speed. (The governor or constant speed unit [CSU] may be an electronic device that detects the rotational speed of a slip-ring incorporated in the propeller hub, and controls operation of a servomotor/leadscrew pitch change actuator in the hub assembly. Or, it may be an hydraulic fly-ball governor attached to the engine, using engine oil to operate a hydraulic pitch change piston in the hub assembly. In the first case, the cockpit control device is likely to be knobs and switches. In the hydraulic system, the governor is likely to be cable operated from a cockpit lever — JB.) A constant-speed propeller will automatically deliver you the advantages outlined above for variable-pitch propellers, with almost no control required from the pilot. Once a propeller/engine speed is selected, the pilot is able to control the power purely with the throttle (actually controlling the absolute pressure of the fuel/air mix in the intake manifold [MAP] which then determines power output) and the controller will act to keep the propeller/engine speed at the selected setting. While allowing the pilot to ignore the propeller for most of the time, the pilot must still choose the most appropriate engine/propeller speed for the different phases of flight: Take-off, go-around and landing. A high speed setting is used when maximum power is needed for a short time such as on take-off. The high speed setting may also be used to keep the propeller pitch low during approach and landing, to provide the desired drag and be ready for a go-around should it be required. Climb and high speed cruise. A medium speed setting is used when high power is needed on a continuous basis, such as during an extended climb, or high speed cruise. Economic cruise. A low speed setting is used for a comfortable cruise with a low engine speed. This operation produces low fuel consumption and longer range, while the advantages of low noise and low engine wear are also enjoyed. When operated in automatic mode, the Airmaster propeller is an example of a constant-speed propeller. Special pitch modes As well as the ability to vary the pitch of the propeller to optimise the aircraft performance, some variable-pitch propellers have some other special modes of operation that can be very useful in certain circumstances: Feather. A feathering propeller can alter the pitch of the blades up to almost 90 degrees. That is, the blade pitch is changed so that they have their leading edge pointing right into the direction of flight, offering minimum resistance to the airflow. This mode allows the propeller rotation to be stopped, without adding excessive drag to the aircraft. Feather may be used to improve the performance of the aircraft after the failure of an engine, but more usually in light aircraft it is used in motor glider applications. Here the engine is used to gain altitude, before the engine is switched off, the propeller feathered, and then gliding flight commenced. Reverse. A reversing pitch propeller can alter the pitch of the blades to a negative angle. That is, the blade pitch is changed so that they have their leading edge pointing slightly opposite to the direction of flight. This mode allows reverse thrust to be developed by the propeller. In larger commuter and transport aircraft this feature is often used to slow the aircraft rapidly after landing, but in sport aircraft it is more usually used to enhance manoeuvring on the ground. A popular application is in seaplanes, where the ability to manoeuvre backwards, and sometimes to reduce the thrust to nothing, is especially useful. This overview was designed to assist the understanding of how the ability to control propeller pitch is used to categorise the different families of propeller design. More importantly it has illustrated that as we progress from one design family to another, we realise significant improvements in performance, effectiveness and efficiency. While a family of propellers that offers better performance is likely to be more expensive to purchase, you can expect that over time the efficiency of a higher performance propeller will produce savings that will offset the initial cost. In addition your flying will be a more relaxed and enjoyable experience! When deciding what type of propeller to buy for your aircraft, you have to weigh up the relative advantages and costs. To help, we can summarise the most common families of propellers, and make a simple comparison of their respective advantages in cost and capability. Marcus Graney Aeronautical Engineer November 2000 There is another type of propeller that is quite rare; the single-blade propeller and, more particularly, the single-blade folding propeller associated with low-power engines in motor-gliders, see 'Single-blade propellers in very light aircraft'. 3.7.4 Propeller theory The forces Propeller blades are constructed using aerofoil sections to produce an aerodynamic force, in a similar manner to a wing. Consequently, the blades are subject to the same aerodynamics — induced drag, parasite drag, wingtip vortices, lift/drag ratios at varying aoa, pressure distribution changing with aoa, etc. There is a difference in application because, in flight, the propeller has rotational velocity added to the forward velocity. Thus, the flight path of any blade section is a spiral — a helical flight path. The diagram at left represents a blade section in flight and rotating about the shaft axis. Because of the different application, it doesn't serve much purpose to express the resultant aerodynamic force as we would for a wing; i.e. with the components acting perpendicular (lift) and parallel (drag) to that helical flight path, as in the upper figure. So, we resolve the aerodynamic force into the component acting forward and aligned with the aircraft's longitudinal axis as the thrust force, and that acting parallel to the direction of rotation as the propeller torque force. As you see in the lower figure the component of the 'lift' acting in the rotational plane has now been added to the 'drag' to produce the 'propeller torque force' vector. The remaining forward-acting portion of 'lift' is then the thrust. That is why propeller efficiency is usually no greater than 80–85%; not all the 'lift' can be used as thrust, and the propeller torque force consumes quite a bit of the shaft horsepower. The propeller torque and the engine torque will be in balance when the engine is operating at constant rpm in flight. Centrifugal force imposes considerable tensile stress while trying to pull the blades from the hub. Torque reaction applies bending stress to the blades in the reverse direction of rotation while the thrust force tends to bend the outer sections of the blades forward. The centrifugal twisting moment tends to twist the blades to a decreased (finer) pitch and the aerodynamic twisting moment (similar to the wing pitching moment) tends to twist the blades to a coarser pitch. The air inflow at the face of the propeller disc also affects propeller dynamics. Blade angle and pitch Although all parts of the propeller, from the hub to the blade tips, have the same forward velocity, the rotational velocity — and thus the helical path of any blade station — will depend on its distance from the hub centre. Consequently, unless adjusted, the angle of attack will vary along the length of the blade. Propellers operate most efficiently when the aoa at each blade station is consistent (and, for propeller efficiency, that giving the best lift/drag ratio) over most of the blade, so a twist is built into the blades to achieve a more or less uniform aoa. The blade angle is the angle the chord line of the aerofoil makes with the propeller's rotational plane and is expressed in degrees. Because of the twist, the blade angle will vary throughout its length. So, normally the standard blade angle is measured at the blade station, 75% of the distance from the hub centre to the blade tip. The angle between the aerofoil chord line and the helical flight path (the relative airflow) at the blade station is the angle of attack and the angle between the helical flight path and the rotational plane is the angle of advance or helix angle. The aoa and helix angle vary with rotational and forward velocity. The basic dimensions of propellers for light aircraft are usually stated in the form of number of blades, and diameter and pitch with values in inches; e.g. 3-blade 64" × 38". The pitch referred to is the geometric pitch that is calculated for any blade station, but usually the station at 75% radius. Geometric pitch = the circumference (2πr) of the propeller disc at the blade station multiplied by the tangent of the blade angle. Thus, it is the distance the propeller — and aircraft — would advance during one revolution of the propeller if the blade section followed a path extrapolated along the blade angle. e.g. For a blade station 24 inches from the hub centre (0.75r) and a 14° blade angle, the circumference = 2 × 3.14 × 24 = 150 inches, and tangent 14° = 0.25. Thus, the geometric pitch is 150 × 0.25 = 38 inches. Propellers are usually designed so that all blade stations have much the same geometric pitch. Designers may establish the ideal pitch of a propeller, which is the theoretical advance per revolution that would cause the blade aerofoil to be at the zero lift aoa; thus, it would generate no thrust and, ignoring drag, is the theoretical maximum achievable aircraft speed. The velocity that the propeller imparts to the air flowing through its disc is the slipstream. Slip used to be described as the difference between the velocity of the air behind the propeller (i.e. accelerated by the propeller) and that of the aircraft. Nowadays, slip has several interpretations, most being aerodynamically unsatisfactory, but you might consider it to be the difference, expressed as a percentage, between the ideal pitch and the advance per revolution when the the propeller is working at maximum efficiency in converting engine power to thrust power. Slip in itself is not a measure of propeller efficiency; as stated previously, propeller efficiency is the ratio of the thrust power (thrust × aircraft velocity) output to the engine power input. Pitch and velocity The performance of aircraft fitted with fixed-pitch or ground-adjustable propellers is very much dependent on the chosen blade angle. Fixed-pitch propellers limit the rpm developed by the engine at low forward velocity, such as occurs during the take-off ground roll; they may also allow the engine rpm to exceed red-line maximum when the load on the engine is reduced, such as occurs in a shallow dive. Fixed-pitch propellers operate at best efficiency at one combination of shaft power and airspeed. Blade angle is usually chosen to produce maximum performance at a particular flight condition, for example: • Vy climb; i.e. a climb propeller • Vc cruise; i.e. a cruise propeller. The climb propeller is usually chosen when the aircraft normally operates from a restricted airfield or in high density altitude conditions. The climb propeller will produce maximum efficiency at full throttle around the best rate of climb airspeed and will perform fairly well at take-off. But during the initial take-off acceleration, even the climb propeller may restrict the engine rpm to less than 75% power. The cruise propeller will achieve maximum efficiency at 75% power at airspeeds around the design cruising speed but aircraft take-off and climb performance will not be the optimum. The cruise propeller usually has a little more pitch than the standard propeller fitted to the aircraft. A high-speed propeller might be fitted when the aircraft is intended to be operating at, or above, rated power for short periods — in speed competition, for example. A variable-pitch constant-speed propeller allows the engine to develop maximum rated power and rpm during the ground roll, and to develop full power throughout its normal rpm range. With a constant-speed propeller, the pilot controls the inlet manifold absolute pressure [MAP] with the throttle lever and the engine rpm with the rpm control lever or knob/switches. (MAP is the pressure of the air/fuel mixture being delivered to the cylinders and is usually measured in inches of mercury [in/Hg] rather than hectopascals. Standard sea-level barometric pressure is 29.92 in/Hg or 1013.2 hPa.) The aircraft flight manual usually provides the pilot with several combinations of rpm/MAP to achieve a particular power setting. For example, in one particular aircraft, the recommended combinations for 65% power at sea-level are 2100 rpm + 26 in/Hg MAP, or 2200 rpm + 25 in/Hg, or 2300 rpm + 24 in/Hg, or 2400 rpm + 23 in/Hg. So, you can use low rpm and high MAP, or high rpm and low MAP, to achieve exactly the same power output. The 2100 rpm/26 in/Hg low rpm/high MAP combination probably gives more efficient cylinder charging and better combustion plus less friction. The high MAP also acts as a cushion in the cylinders, reducing engine stress. Obviously, if a constant-speed propeller is fitted to an aircraft then an intake manifold pressure gauge — marked with the allowable engine operating ranges — must be fitted, otherwise excessive manifold pressure (which raises the cylinder compression pressure) may overstress the engine. Variable-pitch in-flight adjustable propellers also necessitate fitment of a manifold absolute pressure gauge. 3.7.5 The windmilling propeller The angle of attack of a fixed-pitch propeller, and thus its thrust, depends on its pitch, the forward speed of the aircraft and the rotational velocity. Following a non-catastrophic engine failure, the pilot tends to lower the nose so that forward airspeed is maintained while at the same time the rotational velocity of the engine/propeller is winding down. As the forward velocity remains more or less unchanged while the rotational velocity is decreasing, the angle of attack must be continually decreasing. It is possible (depending on the particular PSRU, blade angle etc.) that at some particular rpm, the angle of attack will become negative to the point where the lift component becomes negative (reverses) and the propeller may autorotate; in effect, driving the dead engine as an air pump. This acts as greatly increased aerodynamic drag, which adversely affects the aircraft's L/D ratio and thus glide angles. The parasitic drag (including the 'reversed thrust') is greater than that of a stationary propeller. The engine rotation may cause additional mechanical problems if oil supply is affected. If the forward speed is increased, windmilling will increase. If forward speed is decreased, windmilling will decrease. Thus, the windmilling might be stopped by temporarily reducing airspeed possibly to near stall — so that the reversed thrust is decreased to the point where the engine airpump torque and friction will stop rotation. This is not something that should be attempted without ample height. Should the PSRU fail in flight, the propeller is thereby disconnected from the engine and may 'freewheel' rather than 'windmill'. In the diagram, the upper figure shows the forces associated with a section of a propeller blade operating normally. The lower figure shows the forces and the negative aoa associated with the propeller now windmilling at the same forward velocity. A variable-pitch propeller may have a feathering facility, which turns the blades to the minimum drag position (i.e. the blades are more or less aligned fore and aft) and thus stops windmilling when the engine is no longer producing power. Such a feature is not usually fitted to a single-engine aircraft, but a few powered recreational aircraft are designed with very low parasitic drag plus wide span, high aspect ratio wings that provide L/D ratios around 30:1, and thus have excellent soaring capability. Propeller parasitic drag will have a relatively high effect on the performance of such aircraft so they are usually fitted with a feathering propeller. The image at left is from a FAA Special Airworthiness Information Bulletin (please read) and shows the change in equivalent parasite drag for both a windmilling propeller and a stationary propeller at blade angles from fully flat to feathered. It can be seen that, in this particular case, the windmilling propeller produces more drag than the stationary propeller up to blade angles of 18 degrees or so. It can be inferred from the preceding material that the windmilling vs stationary drag characteristics for aircraft/propeller combinations will be subject to considerable variation. Some motor-gliders are designed with the engine/propeller unit mounted on a retractable pylon, so that when good atmospheric lift conditions exist the engine plus propeller can be stopped and stowed within the fuselage. 3.7.6 The runaway propeller As a propeller system increases in complexity, then the possibilities for malfunction increase. A problem associated with constant-speed propellers is governor failure during flight which, in most installations, will cause the propeller blades to default to their fine pitch limit. This greatly reduces the load on the power plant, and the engine will immediately overspeed, particularly if in a shallow dive. Depending on the fine pitch limit setting, the rpm of an overspeeding engine — sometimes referred to as a 'runaway prop' — may quickly go way past red-line rpm and, unless immediate corrective action is taken, the engine is likely to self-destruct and/or the propeller blades break away from the hub due to the increased centrifugal force. The corrective action is to immediately close the throttle and reduce to minimum flight speed by pulling the nose up. (But see 'Recovery from flight at excessive speed'.) Once everything is settled down, fly slowly, consistent with the fine pitch setting, to a suitable airfield using minimum throttle movements. (The constant-speed propeller fitted to a competition aerobatic aircraft usually defaults to their coarse pitch limit to prevent overspeeding, but an immediate landing is required.) Propeller theory is complex and not appropriate to this Flight Theory guide, but the outline above at least introduces some of the everyday terms encountered. Things that are handy to know The term 'brake horsepower' is a measure of the power delivered at the engine output shaft; measured by means of a dynamometer or similar braking device. The term 'shaft horsepower' [shp] is a measure of the engine power available at the propeller shaft. Generally it is the same as bhp but if the coupling is not direct drive — a propeller speed reduction unit [PSRU] is interposed between the crankshaft output and the propeller shaft as in the Rotax 912 — the shp will be a little less than bhp because of the power loss in driving the belt or gear driven PSRU. The use of the horsepower term for piston aero engines has successfully withstood metrication. To convert horsepower to watts multiply by 745.7 or by 0.75 to convert to kilowatts. When torque is expressed in newton metres, and engine speed in radians per second, power will be in watts. The stoichiometric (chemically correct) air/fuel mixture produces complete combustion of all the fuel and all the oxygen in the cylinder charge — and also the highest temperatures, which may be detrimental to the engine metallurgy. The stoichiometric air/fuel ratio for gasoline fuels is 14.7:1 by weight. Spark ignition engines provide best power with an air deficiency of 5–15% from stoichiometric — i.e. about 12–13:1 (rich) — and provide minimum fuel consumption with around 10% excess air; i.e. about 16:1 (lean). This indicates that the engine, at sea-level and using a stoichiometric mixture, would process about 8500 litres of air per litre of fuel. (Avgas weighs 0.71 kg per litre, and air (at standard sea-level conditions) weighs 1.225 kg per 1000 litres.) The leaned mixture for best economy cruise is around 16:1 (9000 litres of air), and for maximum engine rich mixture performance, around 12:1 (7000 litres of air). The Rotax 912 1.2 litre engine produces 75% power at 5000 rpm, and with a firing cycle every second revolution it would process 1.2 x 5000/2 = 3000 litres of air/fuel mixture per minute. The fuel used would be 3000/9000 = 0.33 litres/minute or around 20 litres/hour, at sea-level. Most four-stroke, normally aspirated, aero-engines between 80 and 400 hp have a specific fuel consumption close to 0.19 kg or 0.27 litres, per horsepower per hour (or 0.42 lbs/hp/hr). Then the Jabiru's engine, rated at 80 hp, but using only 65% for the 97 knot cruise, would consume 80 × 0.65 × 0.27 = 14 litres over 100 air nautical miles, or 7 air nautical miles per litre. Note that you can create a little rule of thumb here that is applicable to most four-stroke engines — "the fuel burn, at 'performance cruise speed', is about one-fifth of the rated engine horsepower — in litres per hour." Thus, fuel burn for the Jabiru cruising at 75% power is 80/5 = 16 litres/hour. Two-stroke engines have to use a richer mixture to run cooler so, for such engines, add about 10% to the calculated result. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.6.1 Lift generation In the 'Basic forces' module it was stated that when an aircraft is moving through the air, the consequent pressure changes or aerodynamic reactions to its motion will be acting at every location on its surface. We had a look at the formula for calculation of lift from the wings: (Equation #1.1) Lift [ newtons] = CL × ½rV² × S It is usual to substitute the symbol 'Q' to represent dynamic pressure [½rV²] so the expression above may be more simply presented as: (Equation #4.1) Lift [newtons] = CL × Q × S where Q × S is a force. It is appropriate to state here that the formula is an approximation of the average lift from the wings. At any one time, the aerodynamic reactions will vary over the span of the wing and with the position at which the wing control surfaces are set. Aerofoils and the aerodynamic force An aerofoil (airfoil, parafoil, wing section or wing profile) is an object — with the shape of the cross-section of the wing — having the function of producing a controllable net aerodynamic force by its motion through the air. To be useful this aerodynamic force must have a lifting component that is much greater than the resistance or drag component. In a powered aircraft, motion through the air is provided by the thrust; so in effect, the aerofoil is a device that converts thrust into lift; in a glider the aerofoil converts much of the gravitational force (the potential energy of height) into lift. The aerodynamic force has two sources: the frictional shear stress, or skin friction, that acts tangential to the surface at every point around the lifting body; and the pressure exerted perpendicular to the surface at every point. (At speeds over about 250 knots, flow compressibility introduces other factors.) The resultant net aerodynamic force is the sum of all those forces as distributed around the body. For wings, it is conventional to show the resultant force as acting from an aerodynamic centre and resolved into two components: that acting perpendicular to the flight path is the lift, and that acting parallel to the flight path is the drag. For propeller blades, the aerodynamic reaction is resolved into the thrust component and the propeller torque component. For rotor blades, a more complex resolution is necessary. Note: normally the aerofoil is incorporated into a wing with upper and lower surfaces enclosing the load bearing structure. However, when designing a low speed minimum aircraft such as the Wheeler Scout there are advantages in using a 'single surface' cambered aerofoil wing, very similar to a hang glider wing. Such wings incorporate a rounded leading edge (formed by the aluminium tubing leading edge main spar) that directs the airflow into the upper and lower streams at all angles of attack. The slight camber is formed by battens sewn into sleeves in the 'sails'. Such wings are somewhere between a thin curved plate and a full aerofoil, and are similar in cross-section to a bird's wing. A parachute wing uses the ram air principle to form the aerofoil shape — see 'The ram-air parachute wing'. Now we need to establish how that airflow actually produces the lifting force. John S Denker has published a web book 'See How it Flies' that has a particularly good section on lift generation with excellent illustrations. You should carefully read through section 3 'Airfoils and airflow' and particularly acquaint yourself with the Eulerian approach of 'streamlines' to visualise airflow. In the illustrative diagram at left, narrowing (A) of streamlines indicates accelerating local speed and decreasing local pressure — a favourable pressure gradient. Opening up (D) of streamlines indicates flow deceleration and increasing pressure — an adverse pressure gradient. The term 'free stream' is usually substituted for 'flight path' when discussing aerofoil characteristics because the aerofoil is presumed stationary, as in a wind-tunnel, and the airstream flows around it. The following summarises the content of section 3 of 'See How it Flies': • A flat plate, held at a small aoa, will generate an aerodynamic force — lift and drag — and indeed, some low momentum aircraft do use basically flat plates as their tailplane surfaces. As mentioned above, the shape of sail-type wings is somewhere between a plate and the more usual wing. However, for aircraft that cruise in the 50–150 knot range, a wing with a rounded leading edge, a sharp or square-cut trailing edge, a cambered upper surface and a flat or slightly cambered bottom surface — i.e. a full aerofoil section — will be far more efficient — aerodynamically and structurally — and more effective in performance. (The faster the aircraft, the more the aerofoil section tends to flatten out. So, for supersonic aircraft we are nearly back to the sharp-edged flat plate.) Aerofoil characteristics The straight line joining the leading edge (left) and trailing edge (right) is the chord line. The curved mean camber line is drawn equidistant between the top and bottom surfaces, and the light coloured gap between the chord and mean camber lines represents the camber — which, in this particular aerofoil [a NACA 4415], equates to 4% of the length of the chord at its maximum point which occurs at 40% of chord length from the leading edge. Aerofoil thickness is the distance between upper and lower surfaces. The maximum thickness of this aerofoil equals 15% of the chord; that is called the 'thickness ratio'. At the trailing edge the included angle between the upper and lower surfaces is significant in wake generation — a lower angle is better, and if the trailing edge is square-cut the thickness there should not exceed 0.5% of the chord. In flight, the angle the wing chord line subtends with the flight path is the geometric angle of attack. • A cambered wing will still produce lift at zero, and slightly negative, geometric angles of attack, as shown in the lift coefficient diagram. The aoa where no lift — only drag — is produced is called the zero-lift aoa which, in the diagram, is nearly –2°. From that diagram you can infer that camber contributes a lift coefficient of about 0.2 and anything greater must be provided by aoa. Of course, this will vary with the amount of camber in a particular aerofoil. If the aoa was reduced below the zero-lift value, for example –4°, then the direction of lift would be reversed. The only time you would need such a negative aoa is when you are flying inverted, or performing aerobatics, neither of which are currently allowable in aircraft registered with the RA-Aus. At the zero-lift aoa, all the aerodynamic force is acting parallel to the free stream and is mostly skin friction drag, with a less significant amount of pressure drag but the latter will increase as the aoa is increased. Pressure drag is explained in section 4.7 'Parasite drag'. Cambered wings perform quite well in inverted flight, but are not as efficient as in normal flight because a higher aoa is needed to make up for the lower wing surface having the maximum camber when inverted. For this reason, aerobatic aircraft tend to use symmetrically shaped aerofoils — i.e. the 'camber' of the bottom surface balances the 'camber' of the top surface and aerodynamically the result is zero camber — thus such wings rely purely on the geometric aoa to produce lift. • At positive angles of attack there is a stagnation point, or line, just under the leading edge of the aerofoil where some of the airflow has been brought to a standstill. The air molecules reaching that line, in the incoming stream, are equally likely to go under or over the wing. Stagnation pressure, the highest in the system, exists along the stagnation line. The location moves down and under the leading edge as aoa increases, up to the stalling aoa. Another more confined stagnation point exists at the trailing edge. If an imaginary line is drawn between the two stagnation points, the cross-sectional view of the division of the aerofoil into upper and lower flow areas becomes apparent. • The behaviour of the airstream flowing around such a wing accords with Bernoulli's principle. As the air accelerates away from the stagnation line, the local airflow over the upper surface gains a greater speed than the lower. Consequently, to retain constancy, the static pressure on the upper surface will decrease, and on the lower surface it may decrease very slightly at low aoa but will increase as aoa increases. There is another concept for explaining the pressure differential between upper and lower wing surfaces. Leonhard Euler was a mathematician who was a contemporary of, and collaborator with, Daniel Bernoulli. The Euler Equations (a special case of Newton's Third Law of Motion) express the relationship between flow velocity and the pressure fields in frictionless flow. Because the air particles follow the curved streamlines above the upper surface, there must be a centripetal force across the streamlines that accelerates the flow towards the centre of curvature. That force must be associated with a pressure gradient across the streamlines; i.e. ambient atmospheric pressure at some distance from the surface, grading to a lower pressure on the upper wing surface. For more information enter the terms 'Euler curvature airfoil OR aerofoil' into a search engine. • The usual way of looking at the lift force is that the wing produces an upflow in the air in front of it and a downwash behind it. That downwash continuously imparts momentum — with a downward velocity component — to the air affected by the passage of the aircraft. As you will recall from the 'Basic forces' module the action of adding downward momentum will have an equal and opposite reaction, which in this case is an upward force applied to the wing. And, of course, the energy provided to impart momentum to the air comes from engine power; in a glider it would come from the gravitational potential energy of height. There is a distinction between the 'downflow' produced by the aerofoil and the additional 'downwash' produced by wing vortices (see below), the deflection of which increases with angle of attack. However, for our purposes we can treat all the momentum imparted to the airstream as 'downwash'. You will also recall, from the 'Basic forces' module, that thrust is the reaction from the momentum imparted to a tube of air with the diameter of the propeller. The associated slipstream or 'prop wash' is the added momentum — quite apparent if you stand behind a stationary aircraft when 'running-up' the engine. Helicopter rotor blades are long, slender rotating wings — somewhere between variable pitch propeller blades and normal wings — and the momentum applied to the air — the 'rotor wash' — can be seen clearly by its effect on dust, vegetation and other objects (like parked ultralights) beneath a hovering helicopter. Similarly, a wing producing lift continuously accelerates a flattened tube of air with diameter approximating the wing span; the longitudinal downward inclination to the flight path of that flat tube increases as aoa increases. Some liken that concept to the wing acting as an airscoop. • Another concept associated with the aerodynamic force — circulation theory — is a mathematical description of a 'bound vortex', which also fits in with the generation of the physical wing-tip vortices. Vorticity is rotary motion in a fluid, and you could regard 'circulation' as referring to the apparent flow rotation — upwash then downwash — around the upper/lower surfaces. Note: there is a long-held and still-continuing argument, particularly in newsgroups and other internet venues, about the pros and cons of the various lift generation theories. None of the arguments put forward (often ill-informed) affect in any way how a light aircraft flies, how it should be safely and economically operated, or how it should be built; so it is best to ignore them unless you are particularly interested in the science of aerodynamics and skilled in mathematics. Pressure differential At any aoa between the zero lift and stalling angles, the total pressure pushing down on the wing upper surface will always be less than the total pressure pushing up on the lower surface. The absolute pressure difference between the upper and lower surfaces will increase as aoa increases up to the stalling aoa. Although it is still small in comparison with the ambient atmospheric pressure, it is this pressure differential resulting from the wing deflecting the air that initiates the lifting force; and this is true however lift theory may be expounded. Much work has been done in designing aerofoils that will maintain the required pressure difference in the targeted flight conditions. We can calculate the net pressure difference for the Jabiru using the scenario in the 'Basic forces' module section 1.4; i.e. cruising at 6500 feet, airspeed 97 knots or 50 m/s, air density 1.0 kg/m³. The ISA atmospheric pressure at 6500 feet is about 800 hPa: static pressure = 800 hPa dynamic pressure = Q = ½rV² = ½ × 1.0 × 50 × 50 = 1250 N/m² = 12.5 hPa Multiplying the dynamic pressure of 1250 N/m² by the lift coefficient of 0.4 gives the pressure differential of 500 N/m². That pressure differential of 500 N/m² (5 hPa) is less than 1% of the ambient static pressure, but applying that over the 8 m² of wing area gives the lift force of 4000 newtons that we calculated in section 1.4. Lift coefficient The lift coefficient CL is a dimensionless (or nondimensional) quantity (it has no units of measure) relating mostly to aoa. It increases as the aoa increases from the normal aoa used in cruise flight, and also to the form of the wing and the aerofoil section. CL represents the proportion of total dynamic pressure converted to lift force. When the aircraft designer calculates the CL curve for an aircraft it must be related to a particular wing reference area. This may be the visible plan area of the wings but it could also include that area of the wings conceptually enclosed within the fuselage. Note that the CL for an aerofoil will have a value perhaps 10–20% higher than the CL for any wing incorporating that aerofoil; this is discussed in the spanwise pressure gradient section. (The convention is to use a lower case 'L' [thus Cl ] when referring to the lift coefficient for an aerofoil to distinguish it from the lift coefficient for a wing, but I have retained CL for both.) In level, non-manoeuvring flight, lift equals weight, so equation 4.1 can be restated as: (Equation #4.2) CL = W / (Q × S) The usable value of CL in a very light aircraft with low-aspect ratio wings without lift-enhancing devices might range between 0.1 and 1.6. (Unless it is a symmetrical aerofoil — same camber top and bottom — the lift coefficient range will be different for the same wing when in inverted flight.) However, a very low CL value can be obtained momentarily if the wings are 'unloaded' in flight. This can be achieved by applying sufficient continuous forward pressure on the control column to attain a near-zero aoa such that the net pressure differential between the upper and lower wing surfaces is very low. This would imply low lift generation and reduced drag, so the thrust will accelerate the aircraft a little faster than normal. Furthermore, a negative CL can be obtained by maintaining so much forward pressure on the control column that the aerodynamic force is reversed. If initially flying straight and level, the aircraft will 'bunt'; i.e. enter the first few degrees of an outside loop with the centripetal force for the turn being supplied by the reversed lift. (This reverses the direction of the wing loading and should never be attempted in weight-shift aircraft nor three-axis aircraft unless the three-axis manufacturer's flight manual allows such a manoeuvre.) And, of course a suitably equipped aircraft can be flown in inverted level flight — in which case the under-wing surface becomes the upper and a completely different CL range applies, because the cambered surface is now underneath and a higher aoa is necessary to maintain the lift required for level flight. Incidentally many pilots utilise the low CL technique when landing a taildragger. The application of forward pressure on the control column after touchdown 'pegs' the aircraft down by reducing the aoa and thus generated lift, and thereby puts increased pressure on the tyres, and amplifies friction and any braking force applied. The same technique was used to bring military DC3 aircraft to a quick stop. 3.6.2 Aerofoil simulation Whichever way lift theory is expounded, this simple equation is applicable: Lift = CL × Q × S I suggest you try out what you have learned so far in an aerofoil flight test simulation program. You need a Java-enabled browser. Read the instructions carefully and reset the measurement units from pounds to newtons. In this case, airspeed will be shown in km/h but just mentally divide by two (and add 10%) to get knots — halve it again if you want m/s. You can try this simple model out with a popular aerofoil, the NACA 2412, which is one of a series dimensioned by the U.S. National Advisory Committee for Aeronautics (the forerunner of NASA) in the 1920s and 1930s. The 2-4-12 (twenty-four twelve) has a camber of 2% [2] of chord with maximum camber occurring at 40% [4] of chord from the leading edge and a thickness/chord ratio of 12% [12]. Note that all dimensions are proportional to the chord so the same aerofoil section shape is retained throughout a wing even if it is tapered in plan form. The wing is thickest at the root and thinnest at the tip; i.e. it must also be tapered in thickness. Most aerofoils suitable for light aircraft have a camber of 2–4%, thickness ratio of 12–15% and the maximum thickness (not camber) occurring at around 30% of chord. Now type the following data into the FoilSim boxes using the 'enter' key or use the sliders: Size: chord 1 m, span 8 m (area 8 m²) Shape: angle (of attack) 2°, camber 2%, thickness 12% Flight test: speed 166 km/h (90 knots), altitude 1947 m (6400 feet) Check the results displayed in the black boxes and in the plots. The static air pressure should be 80.0 kPa (800 hPa) and the lift is 4233 N. If you select 'surface pressure' from the output plots, you will see a plot of the pressure distribution across the chord for the upper (white line) and lower (yellow line) surfaces. Anything appearing above the green line (the atmospheric static pressure) can be regarded as a positive pressure pushing that surface at that point. Anything below the green line is a negative pressure pulling that surface at that point. The area between the two curves represents the magnitude of the differential pressure distribution. The horizontal axis indicates the percentage distance from the mid-chord position. The pressure gradient plot for the upper surface shows a maximum decrease of around 1.5 kPa (15 hPa) close to the leading edge but changing to a slight positive increase in pressure at the trailing edge. The pressure gradient plot for the lower surface shows an increase in pressure under the leading edge, quickly changing to a decreased pressure of a few hPa then back to a positive pressure from mid-chord back. If you press the 'Save Geom' button, a data table will be displayed showing the pressure and local velocity readings at 19 X-Y coordinate positions on both the upper and lower surfaces. If you now select 'surface velocity' for the output plot, you will see a plot of the local velocity distribution across the chord for the upper (white line) and lower (yellow line) surfaces. You can see that the local velocity increases to about 40% above the free stream velocity a very short distance downstream from the leading edge, then it gradually slows until local velocity is less than free stream velocity at the trailing edge. Now change the airspeed to 110 km/h (60 knots) and the aoa to 12°, and look at the surface pressure and surface velocity plots again. Note the big increase in local velocity that is now some 2.5 times the free stream velocity a very short distance downstream from the leading edge. Also note the big increase in the pressure differential and that most (about 70%) is occurring within the first 25% of the chord. You should do a little exploration starting with the aerofoil design, changing just one value at a time and noting the changes in the upper and lower pressure gradients. For instance change the camber from 2 to 4% (i.e. the NACA 4412 aerofoil) and see the lift generated increase to 6369 N with a CL now 0.74. You can do the same with the flight performance items under pilot control — aoa, altitude and airspeed. Of course, FoilSim doesn't provide any information concerning drag generation or pitching moment. 3.6.3 Boundary layer airflow In the following section I use the concept of the airstream flowing over a stationary wing (as in a wind tunnel experiment) rather than the reality of the aircraft moving through stationary air, for easier explanation. The innermost molecules of the moving air come into contact with the solid surface of the wing (and other parts of the aircraft) and are entrapped by the surface structure of the airframe materials. This is called the 'no-slip condition' and is common to all fluid flows. The interaction between those air molecules and the molecules of the solid surface transfers energy and momentum from the air molecules to the solid surface molecules — producing skin friction drag and shear stress that act tangentially to the surface. Those surface-interacting air molecules retreating from the surface consequently carry less momentum than they did on approach. In the very thin viscous sublayer adjacent to the solid surface, these molecules with reduced momentum move randomly into the fluid a small distance from the surface. The streamwise momentum per unit volume of the molecules that have interacted with the surface is less than the momentum a small distance from the surface. The random mixing of the two groups of molecules reduces the streamwise momentum of the molecules that have not directly interacted with the surface. This exchange of momentum between slower and faster molecules is the physical origin of air viscosity (the resistance to flow when a fluid is subject to shear stress) and of that viscous sublayer or boundary layer comprising the region between the wing surface and the unrestrained or inviscid outer stream. The diagram shows the velocity gradient within the boundary layer; the more turbulent the flow, the steeper the gradient and the greater the shear stress and friction. The atmospheric boundary layer is similar but, of course, on a grander scale. Laminar and turbulent flow The thickness of the boundary layer starts at zero at the wing leading edge stagnation point, but will increase (as an increasing number of molecules lose momentum) until a maximum thickness is reached near the trailing edge. The friction between air layers moving at different velocities within the boundary layer is generally weak, so the flow from the stagnation point is initially made up of smooth-flowing stream lines or laminae — laminar boundary layer flow. But on both the wing upper and lower surfaces not far downstream from the leading edge, the laminar flow, less than 1 mm in thickness, usually transitions to a flow with small irregular fluctuations — turbulent boundary layer flow — and continues to increase in thickness by around 1% of the distance travelled to a maximum near the trailing edge of perhaps 10–15 mm for a 1200 mm wing chord. Drag increases as the boundary layer thickens. The extent of laminar flow and thus the location of the transition zone — where boundary flow is a mix of laminar and turbulent — depends on the designed aerofoil shape in profile, the angle of attack, contour variations (ripples, waviness) formed during construction and service, the flexibility of the wing's skin, surface roughness/cleanliness, porosity, and the pressure gradient along the wing chord. In the area where the pressure gradient is favourable (i.e. decreasing, thus the flow is accelerating), laminar flow will tend to continue, though becoming thicker, unless something trips it into the more irregular turbulent boundary layer flow — even paint stripes can trip laminar flow. The laminae nearest the skin move slowly and cohesively, thus minimising skin friction drag. In the turbulent flow boundary layer, the air nearer the wing is moving faster and somewhat chaotically, thus greatly increasing skin friction drag. The transition zone tends to occur a particular distance downstream (for a combination of the preceding factors) rather than a percentage of chord even though the aerofoil might be designed for laminar flow for a particular percentage of chord. The aerofoils used for light aircraft wings have very little laminar flow. But specialised high-speed aerofoils are designed to promote laminar flow over perhaps the first 30–40% of the wing chord by providing a favourable pressure gradient for at least that distance (i.e. maximum thickness at 40–50% of chord) and a properly contoured, very smooth, clean, non-flexing, seamless skin. The latter conditions are also important for minimising the thickness of the turbulent boundary layer flow with consequent reduction in skin friction drag and are achievable in composite construction. Flow separation Generally at lower angles of attack, the boundary layer and the outer stream will separate (break away or detach) from the wing upper surface at the trailing edge or perhaps slightly upstream from the trailing edge, causing a thin trailing wake to form between the outer streams. As aoa increases past perhaps 12°, the boundary layer separation on the wing upper surface might tend to move upstream a little. But at the stalling aoa, separation will suddenly move much further upstream, and a thick turbulent wake will form between the two remnant boundary or shear layers and will be dragged along by the aircraft. The reaction to the wing accelerating and energising that previously stationary air is a sudden deceleration of the aircraft, accompanied by a sudden increase in the magnitude of the nose-down pitching moment. Downwash disappears and the rate of loss of lift will increase rapidly as the aircraft slows. Aerodynamicists devote much effort to controlling and energising the boundary layer flow to delay separation and thus allow flight at lower speeds; for example, see vortex generators. More lift and much less pressure drag is generated in attached turbulent boundary layer flow than in partially separated flow. 3.6.4 Aspect ratio Aspect ratio is the wing span divided by the mean wing chord. An aircraft with a rectangular wing of area 12 m² might have a wing span of 8 m and constant wing chord of 1.5 m. In this case the aspect ratio is 5.33. If the span was 12 m and the chord 1 m, then the aspect ratio would be 12. However because wings have varied plan forms, it is usual to express aspect ratio as: Aspect ratio = wing span² / wing area It is conventional to use the symbol 'b' to represent span, so the equation above is written as: (Equation #4.3) A = b² / S The Jabiru's aspect ratio (span 7.9 m, area 8.0 m²) = 7.9 × 7.9 / 8 = 7.8, whereas an aircraft like the Thruster would have an aspect ratio around 6. Consequently you would expect such an aircraft to induce much more drag at high angles of attack, and thus slow much more rapidly than the Jabiru. And incidently, the mean chord (not the mean aerodynamic chord) of a wing is span/aspect ratio. A high-performance sailplane wing designed for minimum induced drag over the CL range might have a wingspan of 22 m and an aspect ratio of 30, thus a mean chord of 0.7 m. There are a few ultralight aeroplanes, designed to have reasonable soaring capability, that have aspect ratios around 16–18, but most ultralights would have an aspect ratio between 5.5 and 8, and averaging 6.5. General aviation aircraft have an aspect ratio between 7 and 9, probably averaging around 7.5. Note that the higher the aspect ratio in powered aircraft, the more likely is wingtip damage on landing. Note that 'wing area' includes the nominal extension of the wing shape into and through the fuselage. This would appear quite apt for a parasol wing or a high-wing aircraft, but will no doubt seem odd for a mid or low wing. It is just a means for consistent application/comparison between aircraft designs. The span loading is the aircraft weight divided by the wingspan = W/b. The term sometimes refers to the loads applying at specified stations along the span. 3.6.5 Spanwise pressure gradient There is a positive spanwise pressure gradient (the rate of pressure change with distance) on the upper wing surface from the wing tip to the wing root, imparting an inward acceleration to the airflow close to and above the wing. Conversely, at other than a very small aoa, there is a positive underwing pressure gradient from the wing root to the wingtip, and airflow under the wing acquires an outward acceleration. These spanwise (or more correctly semi-spanwise) pressure gradients on the upper and lower surfaces are caused by the higher pressure air from the undersurface revolving around the wingtip into the lower pressure upper surface. This tip effect results in a near total loss of lift at the wingtip because of the reduced pressure differential, with the loss of pressure differential progressively decreasing with distance inboard. Where these two surface airflows with different spanwise velocities recombine past the trailing edge, they initiate a sheet of trailing vortices. These are weakest near the fuselage and strongest at the wingtips, and roll up into two large vortices, centred just inboard and aft of each wingtip. The vortices increase in magnitude as aoa and lift increase, and so increase the vertical component of, and the momentum imparted to, the downwash. As the centre of each vortex is a little inboard of the wingtip, the vortices also have the effect of reducing the effective wing span, the effective wing area and probably the effective aspect ratio. The vortices also affect the air ahead of the aircraft by reducing the magnitude of the upflow in front of the wing and thus modifying (decreasing) the effective wing aoa, with the greatest effect near the wing tip and little effect near the wing root. When a wing is at a low CL aoa the airstream affected by the wing has a slight downward flow. When it is at maximum CL aoa, that airstream has a more substantial downward flow contributed by the vortices. Because of the reduction in the effective aoa, the wing must fly at a greater aoa to achieve the same lift coefficient that a two-dimensional aerofoil will achieve in the laboratory. Also, the wing tip vortices have a decreasing effect with increasing aspect ratio. This is demonstrated in the diagram where there are three (exaggerated) CL and aoa curves plotted. On the left is the laboratory curve for an aerofoil, in the middle the curve for a high aspect ratio wing utilising the same aerofoil and the curve on the right is for a low aspect ratio version. The red horizontal line connects with a particular CL value, say 1.2. The vertical red lines indicate a different aoa for each curve at the same CL, thus the high aspect ratio wing must fly at a higher aoa and the low aspect ratio wing must fly at a still higher aoa for either to achieve CL 1.2. Or to put it another way, at any aoa the wings produce less lift than the laboratory aerofoil. Also apparent from the diagram is that a higher aspect ratio has the effect of a higher rate of lift increase, as aoa increases, than lower aspect ratio wings. A high aspect ratio wing will have a higher CLmax but a lower stalling aoa than a low aspect ratio wing utilising the same aerofoil. Induced drag has a direct relationship to aspect ratio; see section 4.6. Wing-tip vortices make up most of the wake turbulence created by an aircraft in flight and are certainly the most hazardous to following aircraft. They are usually referred to as wake vortices in the context of air traffic and are the same as other atmospheric vortices in that there is a central low pressure core that is often visible as condensation trails when an aircraft pulls higher g in a humid atmosphere. Read the New Zealand Civil Aviation Authorities booklet 'Wake Turbulence'. 3.6.6 Induced drag As explained in section 4.5 the effect of the vortices is to reduce the effective aoa of the wing compared to that of the laboratory aerofoil, which has the further effect of giving a more rearward inclination to the resultant aerodynamic force for the wing, compared to the aerofoil, at a particular geometric aoa. When that aerodynamic force is resolved into lift and drag components, the additional inclination will produce a reduced lift vector (apparent in the preceding CL/aoa diagram) and an increased drag vector. That increase in the drag vector is the induced drag. Induced drag is least at minimum aoa and greatest at maximum aoa. It is often said that the induced drag is the energy dissipated to induce lift; i.e. if CL is increased, induced drag increases, so thrust must be increased to provide additional energy — if the aircraft's flight path is to continue as before. For example, if the pilot wants to increase aoa and maintain the same airspeed (as in a constant rate level turn), then thrust must be increased to counter the increase in induced drag. There is a point in an aircraft's flight envelope where, because of the increasing induced drag, the slower you want to fly the greater the power you must apply — known as 'flying the back of the power curve' — which is opposite to the norm of applying power to fly faster. Elliptical lift force distribution As stated in section 4.5, with most wings — particularly rectangular wings — the higher pressure air underneath the wing flows around the wing tip into the lower pressure area above, thus reducing the pressure differential and the lift; the effect of this decreases as span and/or aspect ratio increase. Induced drag is minimised if the spanwise distribution of the lift forces can be made to present an elliptically shaped pattern, as shown in the diagram, and that aerodynamic load is equally distributed over the wing so that all areas of the wing contribute to load sharing. (This idealised lift force distribution diagram presents a head-on view of the whole wing without any representation of — or distortion by — the fuselage.) . Elliptical spanwise lift distribution will provide a desirable uniform downwash along the span, and can be achieved by choice of wing plan form and/or by twisting the wing to provide something near an elliptical distribution in a speed band selected by the designer. High aspect ratio elliptically shaped (in plan form) wings generally achieve spanwise elliptical lift distribution; however, because of the compound skin curvatures they are the most difficult and time-consuming to construct. Low aspect ratio constant chord (i.e. rectangular) wings without twist are the easiest to construct but generate the most induced drag; however, the introduction of twist makes such a wing much more efficient. Medium aspect ratio wings with a medium taper ratio plus twist are probably the most used shape. Taper ratio is the ratio of the tip chord to the wing root chord. 'Medium taper' would indicate that the tip chord is greater than 50% of the root chord. Sailplane designers have demonstrated that the most effective high aspect ratio wing is one that has a straight (i.e. non-tapered) trailing edge with a leading edge that is increasingly tapered in sections from root to tip. Wing twist or washout The terms 'wing twist' and 'washout' refer to wings designed so that the outboard sections have a lower incidence, 3–4° or so, and thus lower aoa than the inboard sections in all flight conditions. The main reason for wing twist is to reduce induced drag (see section 'Elliptical lift force distribution') and particularly so at a cruising angle of attack or perhaps the climb speed angle of attack. Another reason is to improve the stall characteristics of the wing so that flow separation begins near the wing roots and moves out towards the wingtips. With twist, the sections near the wing root reach the stalling aoa first, thus allowing effective aileron control even as the stall progresses from inboard to outboard. This is usually achieved by building geometric twist into the structure by rotating the trailing edge, so providing a gradual decrease in aoa from root to tip. Washout reduces the total lift capability a little but this disadvantage is more than offset by the wing twist improving elliptical lift distribution and thus decreasing induced drag. Another form of washout — aerodynamic twist — might be attained by using an aerofoil with a higher stalling aoa in the outboard wing sections. Aircraft incorporating washout tend to not drop a wing during an unaccelerated stall. Instead, there is a tendency to just 'mush' down sedately then drop the nose and regain flying speed. The turbulent wake from airflow separation starting at the wing root buffets the tailplane, thus providing some warning of the oncoming stall before it is fully developed. Also, washout is usually applied, for aerodynamic balance, to the swept wings utilised in weight-shift ultralights. However, geometric washout can cause problems at excessive speed. Effect of wing span/aspect ratio on induced drag The equation for calculating induced drag for a wing is: Induced drag = (k × CL² / A) × Q × S where A is the wing aspect ratio [b²/S] and k is related to a span effectiveness ratio. So, induced drag is directly proportional to CL² and inversely proportional to dynamic pressure [Q], and might comprise 50% of total drag at maximum angle of climb speeds. The lower the span loading [W/b](i.e. the greater the physical span or the 'effective' span), the lesser the induced drag at all angles of attack. This results in a decrease in the thrust needed, particularly for climb — or an increase in the potential energy of height for a sailplane. Various wingtip designs, such as Hoerner wingtips, have the effect of moving the vortices slightly further outboard, thereby increasing the effective span and thus reducing the span loading and induced drag. The information in the following box may only be of interest to aircraft homebuilders, so skip it if you wish and go to the next part . Aspect ratio equals b²/S (equation #4.2), so the equation above can be rewritten as: (Equation #4.4) Induced drag = (k × CL² × S / b²) × Q × S The factor k equals 1/Pe where P [pi] equals 3.14 and e is the span effectiveness factor that might vary between 0.7 and 0.9 for the aircraft as a whole. For an elliptic plan form wing, something like that of the near-elliptical wing of the Seafire 46 at left, with (theoretically) no fuselage interference, then e=1.0 and k =1/3.14 × 1.0 = 0.32. A non-twisted tapered wing will have a span effectiveness factor of perhaps 0.9, so induced drag will be 10% greater and greater still (+20%?) for a non-twisted rectangular wing. However, fuselage and fuselage junction interference will reduce the span effectiveness of the wing. Equation #4.2 states that CL = W / (Q × S). Substituting that for CL² in Equation #4.4: Induced drag = k × [W²/ (Q² × S²)] × (S / b²) × Q × S Some of the terms cancel out, leaving: (Equation #4.5) Induced drag = k × W² / (b² × Q) Equation #4.5 shows that induced drag is proportional to span loading squared [W²/b²] and inversely proportional to dynamic pressure [Q], so that two aircraft with quite different aspect ratios but having an identical span effectiveness factor, wing span and weight would produce the same induced drag at the same dynamic pressure (e.g. same density and TAS or lower density and higher TAS, etc). Anything done that gives a small increase in effective wing span will provide a proportionately higher reduction in induced drag. Jabiru induced drag calculation If we guess that the Jabiru aircraft span effectiveness factor is about 0.8, we have enough information to do a rough calculation of the induced drag on our Jabiru cruising at 97 knots at 6500 feet (as in the pressure differential calculation above). We will use a more practical form of induced drag equation for those who skipped the preceding box: Induced drag = k × CL² / A × ½rV² × S For the Jabiru, k = 1/(3.14 × 0.8)= 0.4, aspect ratio [A] is 7.8 and the CL at that speed is 0.4. = 0.4 × (0.4 × 0.4 / 7.8) × (0.5 × 1.0 × 50 × 50) × 8.0 = 0.4 × 0.02 × 1250 × 8 = 80 newtons If you repeat the CL calculation in section 1.4 using the Jabiru's stall speed at 6500 feet, say a TAS of 25 m/s, you will find that CLmax is 1.6. Now if you repeat the induced drag calculations, you will find it has increased fourfold: Induced drag = 0.4 × (1.6 × 1.6 / 7.8) × (0.5 × 1.0 × 25 × 25) × 8.0 = 0.4 × 0.33 × 312.5 × 8 = 330 newtons 3.6.7 Parasite drag Parasite drag is all the air resistance to a light aircraft in flight that is not considered as 'induced', and consists solely of pressure drag and skin friction drag; the latter is due to viscous flow and has been covered in the boundary layer air flow section above. The parasite drag constitutes much of the total aircraft drag at minimum aoa (i.e. high speed) but comparatively little at maximum aoa (minimum speed). Refer to the diagram in section 1.6. When associated with airflow around an aerofoil, the parasite drag is termed profile drag. Pressure drag or form drag is the net pressure differential of those points on the wing; for example, where a component of the pressure acts in the fore and aft direction, and that pressure differential tends to retard the aircraft. Pressure drag, like skin friction, applies to all parts of the aircraft 'wetted' by the airflow. It is greatest for any part of the airframe that presents a flat surface perpendicular to the flow and least for a streamlined shape that has a fineness ratio (i.e. length to breadth) between 3:1 and 4:1. The illustration — a cross-section of a 3:1 fineness ratio wing strut — shows the flow streamlines detaching from the surface close to the trailing edge, with the characteristic wake associated with pressure drag. What is not apparent from the illustration is that, in this instance, the skin friction drag would be significantly greater than the pressure drag There are two specially named classes of parasite drag: interference and cooling drag. Interference drag occurs at the junctions of airframe structures; for example, the junction of the wings and fuselage or the junction of the undercarriage legs and fuselage. The boundary and outer streamflows interfere with each other at the intersections and cause considerable turbulent drag. Interference drag for a well-designed composite aircraft might be 5–10% of total parasite drag but can be very much higher. The cross-flow associated with unbalanced flight (slip/skid) exacerbates interference drag. If interference drag potential is ignored by the designer, vortex development can occur at the wing/fuselage junctions, effectively splitting the spanwise lift distribution into two separate elliptical patterns; this is particularly so with low-wing configurations but not so much with high wings. The problem is minimised, and total parasite drag considerably decreased, by careful design to reduce the number of junctions, and to use fillets and fairing to direct a smooth airflow around the remainder. Usually the most visible evidence of an interference drag reduction program is the large wing root fillet used in low wing aircraft as seen in the AR-5 photograph. Engine cooling drag is normally associated with the cooling airflow for engines enclosed in a drag reducing cowling. The cooling airflow is designed to be efficiently directed from an air intake through a system of baffles for optimum engine cooling, and perhaps to utilise the energy of the added heat to provide a little thrust at the cowling exit point. Where the engine is not cowled, there is a great deal of parasite drag that certainly cools the engine but would not be specially classed as cooling drag. 3.6.8 Aircraft lift/drag ratio In unaccelerated straight and level flight, lift equals weight, and thus will be a constant value. If you look at the total drag diagram in section 1.6 you will see that the drag varies with the airspeed which means, of course, that it varies with angle of attack. The diagram on the left is a plot of the fixed lift value divided by the total drag value; i.e. the L/D ratio, at varying aoa for a reasonably efficient aircraft. It can be seen that L/D [L over D] improves rapidly between zero or negative aoa up to 4–5° then drops off until the stall angle, where the deterioration rate accelerates. Note that a non-aerobatic light aircraft in normal flight would not experience these low L/D values at aoa between 0° and 2°. The maximum L/D for light aeroplanes — a measure of the aerodynamic efficiency of the aircraft — is possibly between 8 and 12. Some of the ultralights designed with wide span, high aspect ratio wings to provide some soaring capability have a maximum L/D around 30. High-performance sailplanes that are built with very wide span, slender, high aspect ratio wings have the greatest L/D, at 40 –50, and thus the greatest efficiency. Powered parachutes have a L/D ratio around 3. There is a limit to the thrust that the engine/propeller can provide (i.e. the drag that it can match) thus there is also a minimum L/D at which maximum engine power is required to maintain constant altitude. Consequently, there will be a minimum aoa (maximum airspeed) and a maximum aoa (minimum airspeed) at which an aircraft can maintain level flight. As there may not be much range between minimum and maximum L/D, the minimum L/D can be quite significant for ultralight aircraft, where a range of engines, some with rather low power, may be utilised in the same model. An under-powered aircraft will perform very badly at the back of the power curve. Glide ratio Maximum L/D usually occurs at an angle of attack between 4° and 5°, or where the CL is around 0.6. This L/D ratio is also termed the glide ratio because it is just about the same ratio as distance covered/height lost in an engine-off glide. For example, if maximum L/D =12 then the glide ratio is 12:1, meaning the aircraft will glide a distance of 12 000 feet for each 1000 feet of height lost, in still air. We can use the '1-in-60' rule to calculate the angle of the glide path relative to the ground; for example: L/D = 12, then 60/12 = 5° glide path angle. If the aircraft is maintained in a glide at a degraded L/D, then the glide path will be steeper: L/D = 8, then 60/8 = 7.5° glide path angle. This is one effect of using flaps (see section 4.11). Be aware that quoted L/D ratios rarely take into account the considerable drag generated by a windmilling propeller. The aoa associated with maximum L/D decides the best engine-off glide speed [Vbg] for distance and the best speed for range [Vbr] according to the operating weight of the aircraft. But because of the flat shape of the curve around maximum L/D, these speeds are more akin to a small range of speeds rather than one particular speed. 3.6.9 Pitching moment When using the FoilSim aerofoil flight test simulation program, the static pressures around the aerofoil are given in the output plot that shows the pressure distribution pattern changing with the aoa. It is convenient to sum that distribution and represent it as one lift force vector acting from the centre of pressure [cp] of the aerofoil or wing for each aoa; much the same way as we sum the distribution of aircraft mass and represent it as one force acting through the centre of gravity. The plot on the left is a representation of the changing wing centre of pressure position with aoa. The cp position is measured as the distance from the leading edge expressed as a percentage of the chord. (Please note the diagram is not a representation of the pitching moment.) At small aoa (high cruise speed) the cp is located around 50% chord. As aoa increases (speed decreases) cp moves forward reaching its furthest forward position around 30% chord at 10–12° aoa, which is usually around the aoa for Vx, the best angle of climb speed. With further aoa increases, the cp now moves rearward; the rate of movement accelerates as the stalling aoa, about 16°, is passed. Most normal flight operations are conducted at angles between 3° and 12°, thus the cp is normally positioned between 30% and 40% of chord. The movement of the cp of the lift force changes the pitching moment of the wing, a rotational force applied about some reference point — the leading or trailing edges for example — which, in isolation, would result in a rotation about the aircraft's lateral axis. The consequence of the rotation is a further change in aoa and cp movement that, depending on the cp starting position may increase or decrease the rotation. Thus a wing by itself is inherently unstable and will change the aircraft's attitude in pitch — i.e. the aircraft's nose will rotate up or down about its lateral axis, which may be reinforced or countered by the action of the lift/weight couple — so there must be a reacting moment/balancing force built into the system provided by the horizontal stabiliser and its adjustable control surfaces. This will be discussed further in the Stability and Control modules. Aerodynamic centre There is a point on the wing's mean aerodynamic chord (see below) called the aerodynamic centre [ac] where the pitching moment coefficient [ Cmac ] about that point is small — for the NACA 2412 aerofoil Cmac is –0.1. The negative value indicates the moment produces a nose-down torque, which is the norm for cambered wings. Cmac remains more or less constant with aoa changes but becomes more nose-down at the stall. For the cambered aerofoils used in most light aircraft wings, that aerodynamic centre will be located in a position between 23% and 27% of the chord length aft of the leading edge, but for standardisation, aerodynamicists generally establish the lift, drag and pitching moment coefficients at the 25% (quarter) chord position. The notation for the pitching moment at quarter chord is Mc/4. The pitching moment is consistently nose-down, changing in magnitude as airspeed changes. When plotted on an aerofoil wind tunnel data graph, the moment coefficient Cmc/4 is a roughly horizontal line for most of the angle of attack range, but the straight line may have a slight slope if the actual aerodynamic centre varies a little from the 25% chord location. Pitching moment equation: (Equation #4.6) Pitching moment [ Mc/4 ] = Cmc/4 × ½rV² × S × c The pitching moment equation is much the same as the lift and drag equations with the addition of the mean aerodynamic chord [c] for the moment arm; using SI units the result is in N·m. As the coefficient is always negative and nearly constant (up to the stall), then V² is the significant contributor to the nose-down pitching torque, which must be offset by tailplane forces to keep the aircraft in balanced flight. However, high torsion loads may still exist within the wing structure; see aerodynamic effects of flight at excessive speed. The concept of the aerodynamic centre is useful to designer/builders, because it means the centre of application of lift can be assumed fixed at 25% chord and only the lift force changes. For non-rectangular wings, a mean aerodynamic chord [MAC] for the wing has to be calculated; see ascertaining mean aerodynamic chord graphically — in that diagram the aerodynamic centre position [ac] is shown on the root chord line. Neutral point It is not just the wings that produce lift, the tailplane surfaces also produce lift (which is discussed in module 6), and so do parts of a well-designed fuselage. Consequently the aerodynamic centre for the aircraft as a whole, known as the neutral point, will not be in the same location as the wing aerodynamic centre but — for a tailplane aircraft — behind it and on the fuselage centreline. This is the fixed point from which net lift, drag and aircraft pitching moment are assumed to act. 3.6.10 Ailerons We mentioned in section 1.4 that the pilot cannot change the shape of the wing aerofoil. But this, like many statements made regarding aeronautics, needs qualification. In fact, the pilot manoeuvres the aircraft in the lateral plane by altering the effective camber of the outboard sections of the wings. And remember in the last paragraphs of section 4.1 above, using FoilSim, we found that altering camber from 2% to 4% produced a substantial increase in CL and lift. If you examine the Seafire photograph, in section 4.6, you will see that each wing has a separated section at the outboard trailing edge. These are ailerons, hinged to the main wing so that they can move down or up and linked, via control rods or cables, to left/right movement of the pilot's control column. The control column is a simple lever which amplifies forces applied by the pilot. Thus the pilot can, in effect, increase or decrease the camber of the outer portion of each wing; as shown by the effective chord lines in figures A and B at left. The ailerons are interconnected so that downward movement — a camber increase — in one is combined with an upward movement — a camber 'reflex' — in the other. The aileron movement then increases the lift generated by the outer section of one wing whilst decreasing that from the other, thus the changed lift forces (at a distance from the aircraft's longitudinal axis) impart a rolling moment in the lateral plane about that axis. This rolling moment is primarily used to initiate a turn but other manoeuvres depend on the amount and timing of aileron movement; more about this in the 'Control' module; see 'Control in a turn'. Ailerons span perhaps the outer 35% of each wing and occupy perhaps the aft 20% of the wing chord at that location. High-speed aircraft may have two sets: a normal outer wing set used only for low-speed flight (because of the moment of force they are capable of applying at high speed) and a second, high-speed set of spoiler-type ailerons located at the inboard end of the wing. Aileron drag Increasing camber and thus CL also increases induced drag (in proportion to CL²) so that the wing that is producing greater lift will also be producing greater induced drag, tending to rotate (yaw) the aircraft's nose in the direction of the lowered aileron. Parasite drag will be increased on the wing with the lowered aileron. This induced plus parasite drag reaction is called aileron drag and particularly complicates aileron effects at low speeds when CL is high, the aerodynamic pressure on control surfaces is low, and it is easy to impart an excessive control movement. Because the yaw is towards the lowered aileron and thus opposite to the required direction of turn, the effect is called adverse yaw and is particularly evident in aircraft that have long-span wings where the ailerons have a much longer moment arm. Aileron drag can have an opposite yaw effect. When an aircraft is turning at low speed and the pilot applies aileron to roll upright, the downwards movement of the aileron on the lower wing might take the aoa, on that part of the wing, past the critical aoa. Thus that section of wing — rather than increasing lift and making the wing rise — will stall and lose lift. The aircraft, instead of straightening up, will roll into a steeper bank. Although the wing section may be stalled, CL and thus induced drag will still be fairly high, so there will be a substantial yaw toward the lower wing which pulls the nose down and increases the rate of descent. There is potential for other aileron-induced problems when turning at low speeds; see 'Control in a turn'. There are a number of configurations which, used singly or jointly, reduce aileron drag. For example, differential ailerons, where the down-going aileron moves through a smaller angle than the up-going aileron or Frise ailerons, where the leading edge of the up-going aileron protrudes below the wing undersurface, increasing parasite drag on the down-going wing. 3.6.11 Flaps The other camber increasing devices, forming part of the inboard wing trailing edge in the Seafire photo, are the flaps. Plain flaps are also a hinged section of the wing — as in figures C and D in the aileron diagram above — but move only (and jointly) downward usually to fixed predetermined positions, each position providing varying degrees of increased lift coefficient and increased drag coefficient that the designer thought appropriate. For instance, for one particular aircraft, at 5° deflection there is a good increase in CL with only slight increase in drag. At 15° the drag increase starts to equate with the increase in the CL, whereas at 25° or 30° the increase in drag is much greater than the increase in CL; at 45° the flap is starting to act as an airbrake. The change in camber (over perhaps 50–60% of the wing span and 20–25% of the wing chord) caused by lowering flaps in flight, without changing other control positions, has effects which will vary according to the amount of deflection employed: The aircraft's nose will pitch down a few degrees about its lateral axis (i.e. its attitude in pitch is altered) because of the nose-down pitching moment associated with flaps. The position of the aircraft's line of drag will change and this also tends to change the aircraft's attitude in pitch. Depending on the relative mounting of the aircraft's wings and tailplane, the change of direction (and the increase) of downwash may affect the trim of the aircraft — nose up or down. The lift increases and the aircraft will initially tend to rise. The drag increases and the aircraft slows below its trimmed airspeed, lift reduces, and the aircraft sinks unless power is increased. The pilot has to take appropriate control action depending on the reason for lowering flaps. The effects of trim associated with lowering or raising flaps for a particular aircraft type will be noted in the Pilot's Operating Handbook. As we saw in FoilSim, the effect of increasing camber is an increase in CL (the ratio of lift to dynamic pressure or airspeed) at all aoa. This is shown in the plot at the left. At an aoa of 6° CL is about 1.0 with flaps lowered — about 50% greater than the CL of 0.65 with flaps raised. What this means is that the minimum controllable flight speed is lower with flaps deployed. So, returning to the equation: lift = CL × ½rV² × S thus for lift to remain constant if CL increases then V² must decrease. Consequently, the stall speed is also lower with flaps deployed. (Incidently, this diagram shows that the zero lift aoa for this wing occurs at –2°.) Note that the flapped section will stall at a lower aoa than the unflapped section. Generally the flapped wing area, being the inboard section of the wing, represents a very large proportion of the total wing area — check the Seafire photo. Also, even if the flapped section has passed its stalling angle, it is still producing lots of lift. Providing there is sufficient thrust available to overcome the big increase in drag, the aircraft can still maintain height and stability because the wing outboard section and ailerons are not stalled. Bear in mind that to maintain the same airspeed and altitude after lowering flaps, that thrust, if available, must be increased to counter the additional drag from the lowered flaps. Similarly, when flaps are raised, the aircraft will initially sink due to the loss of lift unless the pilot takes compensating control action; this is particularly important when a landing approach is discontinued and a go-around initiated. Now what aoa are we measuring? If you look at figure C (in the drawing in section 4.10) which represents the unflapped part of the wing, you can see that it has an aoa of about 5° or so whereas, at the same time, the flap extended section of wing (figure D) has a considerably greater aoa. As the flapped section will still have a stalling aoa around 16° we can surmise that this flapped wing section is going to stall when the unflapped section is only at 13° or so. (The horizontal axis of the plot shows only the aoa of the unflapped wing.) However, we also have to take into account the increased downwash and thus the change in effective aoa associated with it, so the effect of flaps is not as straight-forward as implied in the preceding. Flap systems There are a many types of flap systems, but if flaps are used at all in ultralights or other very light aircraft, then only the simpler devices shown at left are needed. The most common (because of its simplicity) is the plain flap, which might provide a 0.5 increase in CLmax with a large increase in drag when fully deflected. The split flap provides slightly more increase in lift but a larger increase in drag, and is more difficult to construct and thus probably not worth the effort. The slot incorporated into the junction between the main wing and the plain flap in the slotted flap arrangement allows airflow from under the wing to energise (i.e. accelerate and smooth) the turbulent boundary layer flow over the upper surface of the lowered flap. This provides better downstream boundary layer adherence, and thus allows a larger angle of attack to be achieved before stall, with higher CL and lower drag than the plain flap. Ailerons may also be 'slotted' for improved performance. The rearward extension of the Fowler flap as it is deflected increases wing area as well as camber, so it provides the best increase in lift of all the simpler systems — although perhaps even a single-element Fowler flap like that shown is not that simple to construct. Summary — flap effect on coefficient of lift In the diagram above, it can be seen that the deflection of flaps provides an increase in CL of about 0.4 at all angles of attack. This is probably representative of plain flaps extending along 50% of the wing trailing edge with chord equivalent to about 20% of the wing chord, and deflected 25°. The attainable CL increase depends on flap span, chord and degrees deflected, plus the complexity of the flap system — CL increase of 0.8 might be achieved with long-span Fowler flaps deflected to 35°. Incorporating slots into plain or Fowler flaps increases CL. Advantages of using flaps If flaps are fitted, a small flap deflection — say 10° — might be used for safer take-off, due to the lower lift-off speed available. But half to full flap deflection is always used for landing to provide: lower safe approach and touch-down speeds a nose-down attitude for a better view of the landing area a steeper approach path (because of the degraded L/D) for better obstacle clearance, which can be controlled at will a shorter 'float' after rounding out because of increased drag a shorter ground roll, if flaps are left fully extended until the aircraft has exited the runway. And flaps enable the approach to be made with engine power well above idle, which is beneficial to the engine, allows power changes to either increase or decrease the rate of sink and provides better engine response in case of a go-around. Flaperons In some light aircraft designs, particularly those with short take-off and landing [STOL] capability, it has been found expedient to incorporate the aileron and a plain flap into one control surface that extends the full length of the wing trailing edge. The different functional movements are sorted out by a control mixer mechanism. Usually, the flaperon is not integral with the wing but bracketed to the underwing to provide a slotted flap — acting like an external aerofoil flying in close formation with the main wing. Although the CL increase attainable might be 1.0, there are drawbacks to this arrangement, which particularly exacerbate low speed aileron drag. Reflex flaps Some aircraft are fitted with flaps that also can be deflected upward 5° or 10° above the normal neutral or stowed position in addition to the normal downward deflection positions described above. Upward deflection of flaps is done at cruising speed, and increases the maximum cruise speed perhaps 5% by reflexing camber and reducing drag, and is often associated with aerofoils that have good laminar flow. 3.6.12 High-lift devices Another short take-off and landing [STOL] device used in light aircraft is an aerofoil section — a slat — fixed to the leading edge of the wing, with a slot between the slat and the wing. The slat/slot works in much the same way as the slotted flap except that leading edge slats induce a nose-up pitching moment. At low aoa, the fixed slat has no value; it just increases drag and thus degrades cruise performance. At high aoa, the higher pressure on the underside of the slat is channelled through the slot, gaining velocity and energising the boundary layer flow over the upper surface of the wing — thus delaying boundary layer separation, adding perhaps a 0.6 CL increase and increasing the stalling aoa to perhaps 20°. The usual increase in CL and the stalling aoa is illustrated with the green curves in the CL/aoa diagram above. Some slat/slot systems also have the effect of increasing wing area thus reducing W/S and stall speed. Leading edge slots combined with long-span slotted flaps, as used in STOL aircraft, allow a critical aoa much greater than the usual 16°. They can perhaps double the maximum CL of the basic wing, which allows much lower landing speeds but requires flight at the back of the power curve. Fixed leading edge slots work particularly well with a tailwheel configuration in a 'utility' aircraft such as the Slepcev Storch, but in a touring aircraft they have no value unless the pilot intends operating into very small, rough airstrips. There are simple automatic slat/slot systems where the slat is stowed when flying at lower angles of attack but pops out to form the slot when a particular angle of attack is reached. There are also retractable slat/slot systems that provide STOL capability when required without sacrificing cruise performance, except for the weight increase due to the more complex operating system. I suggest now you have a look at the diagrams in Anatomy of a STOL aircraft. 3.6.13 Lift spoilers and airbrakes The converse of the high-lift devices is the light aircraft spoiler, common in gliders but occasionally seen in high L/D ratio ultralights. The usual spoiler is a flush-mounted front-hinged spring-loaded flat plate incorporated into the upper wing surface, which can be elevated by lever operation to varying degrees of opening. When activated, it induces separation over part of the wing, thereby acting as a lift-dumper. But it is not speed limiting; the nose will pitch down and the pilot must use elevator to maintain the required approach speed; thus the spoiler is used to increase the sink rate on the approach path. Airbrakes or speedbrakes have a similar but more effective function. They are often vertically mounted plates, pairs of which are incorporated into the wing structure and which protrude from the upper and lower wing surfaces when activated. They create a lot of drag but little or no change in pitch, so the pilot must lower the nose to maintain approach speed. Airbrake or spoiler configurations are sometimes associated with flap systems that are primarily directed to lift generation, rather than lift generation plus drag creation. Such flap systems would have maximum downward deflection of perhaps 20°. Military aircraft utilise very complex flaperon/spoileron systems. Things that are handy to know The aerofoil is often referred to as a 'two-dimensional' object. This means that that the spanwise — thus 'third-dimensional' — pressure gradient effects associated with a normal wing, and varying significantly with the wing form rather than the aerofoil shape, are ignored when considering aerofoil characteristics. Wing upflow: all the air disturbances caused by the passage of an aircraft are propagated as pressure pulses moving outward (from molecule to molecule) in all directions at the speed of sound. Thus, in subsonic flight, the pressure variances (compression then relaxation) contribute to the air upflow occurring in front of the wings. In sport and recreational aviation the term aircraft is a generic covering all types of aerial (airborne) vehicles; it includes 'lighter-than-air' (aerostats) and 'heavier-than-air' (aerodynes) but not vehicles that derive their lift from air reaction with the surface, e.g. hovercraft. The aerostats include hot-air balloons and power-driven hot-air airships, both deriving lift from buoyancy. The aerodynes derive their lift from the aerodynamic reactions described above and are in two classes — rotary-wing (rotorcraft) and fixed-wing. Rotorcraft are represented by helicopters, gyroplanes and the towed gyrogliders or rotor-kites. The fixed-wing aerodynes may be power-driven or unpowered, the latter represented by the various glider classes — sailplanes, hang gliders, paragliders and the towed parasails or para-kites. The power-driven aerodynes are represented by three groups: the weight-shift controlled trikes, powered parachutes, powered hang gliders and powered paragliders. the 3-axis controlled power-assisted sailplanes and motor-gliders and finally the ubiquitous 3-axis controlled aeroplanes. For more information see sport and recreational aircraft categories. Notes for homebuilders • The parasite drag coefficient. The equation for calculation of the total parasite drag for an aircraft is: Parasite drag [newtons] = CDp × ½rV² × S Unlike the lift coefficient, the parasite drag coefficient CDp is more or less a constant — the ratio of drag to dynamic pressure — and thus provides a means for comparing the relative aerodynamic 'cleanness' of two aircraft. The coefficient is usually in the range 0.03 to 0.08 for fixed-undercarriage aircraft. • There is another value, the 'equivalent flat plate area' [FPA] used by aircraft, motor vehicle and structural engineers who are concerned with the calculation of air resistance. FPA is often quoted in aviation magazines when comparing the parasite drag efficiency of an aircraft with other similar aircraft, and it is usually stated in terms of square feet. FPA is calculated as CDp times the wing area divided by the CDp for a flat plate. However, it is assumed that the CDp for a flat plate held at 90° to the airstream = 1 (in fact it is about 20% greater, but that is of no real consequence) so the flat plate CDp is omitted from the calculation, thus: FPA = CDp × S ft² For example, the FPA for the run-of-the-mill two or four-seater fixed-undercarriage general aviation aircraft would be around 6 ft² with CDp of 0.03 to 0.05, and the retractables around 4–5 ft² with CDp of 0.02 to 0.03. FPA of a very clean, high-performance general aviation aircraft like a Mooney model, is around 3 ft² with CDp about 0.015. Some very clean, high-performance GA kit-built aircraft have FPA less than 2. Note that FPA does not represent the frontal cross-section area of the aircraft. One of the smallest known FPA is not associated with a general aviation aircraft but with an owner-designed and built ultralight! Californian Mike Arnold's 65 hp two-stroke Rotax 582 powered AR-5 held the world speed record, in the under 300 kg FAI efficiency Class C1-A/0 of 213 mph in August 1992. This handsome little glass-epoxy aircraft has an FPA of 0.88 ft² with CDp about 0.016. It demonstrates the efficiency that can be achieved — an unmatchable 3.3 mph per hp — in an ultralight design when the home designer/builder pays the utmost attention to detail. Note the drag reduction achieved by the beautifully shaped engine cowling, the wing root fillet and the minimisation of the junctions of undercarriage leg fairing and wheel cover. The choice of a fibreglass/foam lay-up composite structure also facilitates the drag reduction program. Don't let anyone tell you ultralights have to be slow and draggy! • 'Separation bubbles' or 'laminar flow separation bubbles'. In laminar flow, sometimes the laminar flow boundary layer separates from the wing surface then reattaches itself a short distance downstream. This forms a 'bubble' of stagnant air with a significant spanwise dimension that changes the aerodynamic thickness of the wing which, in effect, increases pressure drag. Bubbles may also cause increased turbulent flow to be generated downstream of the reattachment point. Aircraft designers avoid laminar separation at cruising speeds by inducing a turbulent — but attached — boundary layer where necessary. Separation bubbles that increase drag may also occur on the fuselage and tailplane. See vortex generators below. • Reynolds number. Occasionally, in reference to boundary layer laminar to turbulent flow and flow separation characteristics you may see mention of the critical Reynolds number [Re]. Re is a measure of the relative influence of viscous and inertia effects on boundary layer behaviour. Like aspect ratio Re is a dimensionless quantity, i.e. it has no units of measurement. For rough estimates Re in airstreams = air density/air viscosity × airflow velocity × flow distance. ISA sea-level density is 1.225 kg/m³ and standard viscosity is 0.0000179 kg/m/s, so standard air density/air viscosity = 68 459 (say 70 000). If the Re is estimated for an average flow across the entire wing at a particular airspeed, then the equation can be simplified to: Re = velocity (m/s) × the mean aerodynamic chord (m) × 70 000. Thus, the wing chord Reynolds number for an aircraft with a MAC of 1.2 m flying at 97 knots (50 m/s) is roughly 50 × 1.2 × 70 000 = 4 200 000. When that same aircraft is cruising at 78 knots (40 m/s), Re would be about 3 360 000. For a particular wing and wing surface condition, there is a critical boundary layer Re, above which the laminar flow will transition to turbulent flow. In slow flight speed, the critical boundary layer Re will be attained a particular distance downstream from the leading edge stagnation point; as airspeed increases (and in accordance with the equation above), that distance must shorten. • The vortex generators [VGs] used in a few light aircraft designs, particularly short take-off and landing [STOL] aircraft — or as post-delivery 'add-ons' — are small boundary layer control devices with a swept-back leading edge and a vertical trailing edge, the chord at the base of the device is two to four times its height. VGs are machined from an aluminium 'T' extrusion or formed in polycarbonate, a row of 10 or 15 of which are usually spaced along the upper surface of each wing, probably close to the transition zone. Each VG is carefully sited, with a specific angle of attack (perhaps 15° or more) to the local airflow and with sufficient height to just intrude into the free-stream flow. So situated, they induce fast-rotating, highly organised, downstream vortices (much the same principle as wingtip vortices) which mix the high-speed free-stream airflow into the slow-moving, surface boundary layer flow; entraining and re-energising that flow so that the chordwise pressure gradient profile on the upper surface is decreased (see boundary layer flow). Consequently, surface pressure is decreased, so the pressure differential — and thus the lift coefficient — is increased at all wing aoas. VGs are often paired, to produce counter-rotating vortices. The VGs also delay boundary layer flow separation at high aoa; i.e. VGs lower the stall speed while improving the aircraft's low-speed behaviour. But there is likely to be minimum warning of onset of the stall, and stall behaviour may be more violent. Appropriately sized and sited wing vortex generators can be effective at providing good manoeuvring control of the aircraft when operating low and slow, and provide a greater CLmax, and improve aileron performance and aircraft climb performance. They are sometimes also used on the horizontal and vertical stabilisers mounted just forward of the rudder/elevator hinge lines where they have the effect of allowing greater control surface deflection before separation occurs. VGs are also useful in locations where interference drag is a problem. The use of VGs in light aircraft may slightly degrade performance at the upper end of the speed range, probably depending on the amount of additional turbulence they generate outside the normal turbulent boundary flow. • The term burble is sometimes used to describe a turbulent stream. For example a disturbance emanating from something on the fuselage can induce a turbulent streamflow that affects the tailplane. There may also be a separation of flow at the junctions of structural components, which causes interference drag. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.5.1 Altitude sensing instruments The sensitive altimeter The altimeter is the cockpit instrument that indicates the aircraft's altitude. The instrument is a refined aneroid barometer with a dial indicating height above mean sea level rather than atmospheric pressure. The main component of such an instrument is a small, flexible, corrugated metal capsule from which the air has been partially evacuated — fitted with a metal closure or diaphragm. There is a spring within the capsule that applies a constant force to the bottom of the diaphragm, while atmospheric static pressure applies a counter force to the top, so that the diaphragm moves as atmospheric pressure changes. The movement of the pressure-sensing capsule is transferred and magnified — via a mechanical linkage or piezo-quartz component — to a dial pointer or pointers, or a digital display, which indicate the altitude reading. The static pressure is drawn from the aircraft's static vent/s, which may induce slight position errors due to aerodynamic effects around a vent. There may be two static vents in different locations on the airframe and the pilot may be provided with the ability to select either or both. The level in the atmosphere at which any particular pressure occurs is also dependent on temperature — as we saw in the 'Airspeed and the properties of air' module — but the altimeter does not sense the air temperature. Consequently, all altimeters are calibrated in accordance with the International Standard Atmosphere [ISA] model, which utilises a standard temperature lapse rate with height of 6.5 °C per km (2 °C per 1000 feet). The atmosphere in any region rarely corresponds to the ISA model, so aneroid altimeters do not indicate totally accurate height. This is not that important, as true altitude can be calculated, in the rare circumstance that it is needed for terrain clearance purposes by an aircraft operating under the visual flight rules. There is no problem with air traffic management, in that all aircraft in the same region, with properly set (and functioning) altimeters, will be out by the same amount. It is, of course, desirable to set the current local surface pressure into the altimeter by setting that reference pressure into a baro-setting scale or 'sub-scale' (known since the 1930s as the 'Kollsman* window'), which in turn resets the position of the height-indicating pointers against the dial. Or, if the aircraft is on the ground, the same result is achieved by turning the baro-setting knob until the altimeter indicates the known airfield elevation. The sensitive altimeter in the image indicates an altitude of about 1410 feet with the baro-scale setting at 29.9 inches of mercury [in.Hg] — equivalent to 1012.5 hPa. If the altitude was 11400 feet, the pointer with the inverted triangle on the end would be past the figure 1 on the image, indicating +10 000 feet. *Paul Kollsman invented his 'sensitive altimeter' in 1929 which was a far superior instrument to those existing at the time but it didn't gain widespread use until 'instrument flying' became common later in the 1930s. In Australia, all barometric pressures are reported in hectopascals (equivalent to millibars); and in the USA in units of inches of mercury (one in.Hg = 33.865 hPa so 29.92 in.Hg = 1013.25 hPa). The baro-scale setting range provided in modern altimeters may be from 800 to 1050 hPa. Electronic altimeter Electronic flight instrument systems [EFIS] use solid-state electronic componentry plus software to display the usual flight instrument readings on a liquid crystal, or similar, screen. In such systems, the atmospheric static pressure is fed to a pressure transducer, which senses and convert pressures to voltages. See the screen display of the Dynon D10A light aircraft EFIS. Note that the EFIS has an outside air temperature probe and the software can calculate density altitude (see section 'Altitude and Q-code definitions') when needed. Electronic altimeters are also available as single instruments or possibly combined with an ASI function. Altitude encoding devices Altitude encoding devices continually supply pressure altitude data (in Gillham 'Gray' code format) to aircraft transponders and/or GNSS receivers – 'baro-aiding'. There are two types; encoding altimeters and blind encoders; the latter are stand-alone digital devices with no display (hence 'blind') probably with a pressure transducer connected to the aircraft's static pressure system. Standard pressure (1013.25) is factory pre-set as the scale basis in all altitude encoding devices so both types send pressure altitude not altimeter-indicated altitude. This pressure setting within the device cannot be altered by pilots, such devices being primarily an air traffic management aid. A user's manual for the Australian Microair EC2002 low power encoder may be downloaded from the Microair website. 3.5.2 Altitude and Q-code definitions Altitude - the third positional dimension An aircraft's 3-dimensionsal position may be very accurately defined by its latitude, longitude and altitude; and the latter is normally the most safety-critical dimension. Contour lines and spot points on WACs and VNCs provide an indication of terrain elevation, i.e. height above the reference datum, which is the Australian Height Datum (AHD). The aircraft's altimeter reading provides the aircraft's vertical position and thus an indication of the current height above the terrain indicated on the chart — height above ground level (agl) or airfield level (afl) and the terrain clearance — may be determined. However, in aviation, that altitude reading and the altitude term itself, have many connotations; particularly important is the concept of density altitude. Altimeter indicated altitude: is the approximate height of the aircraft above the AHD or above mean sea-level [amsl], calculated in accordance with the ISA but using a local or area QNH as the pressure setting rather than the ISA Standard Pressure of 1013.2 hPa. In Australia the AHD represents mean sea level. However an aircraft maintaining a constant altitude — with 1013 hPa or a local/area QNH set in the baro-setting window — is following an isobaric or contour surface whose height above the AHD will vary according to atmospheric temperature and pressure conditions. In the Australian summer temperatures the 'thickness' of the atmosphere is greater than the ISA standard and consequently the rate of pressure decrease with height is less than ISA and the altimeter indicated altitude will be lower than the true altitude. If the atmosphere is colder than the ISA the thickness of the atmosphere is less than the ISA standard and consequently the rate of pressure decrease with height is greater than ISA and the altimeter indicated altitude will be higher than the true altitude. Also note that; if you fly from an area of higher pressure to lower and do not obtain and reset the new area/local QNH the altimeter will be over-reading (the aircraft is lower than indicated) conversely flying from an area of lower pressure to higher the altimeter will be under-reading (the aircraft is higher than indicated) but if you fly from an area that is warmer to a cooler one, the altimeter will be over-reading and conversely, flying from an area that is cooler to one that is warmer the altimeter will be under-reading and the aircraft is higher than indicated. So the adage "From high to low, look out below" is incomplete and the adage "From high to low, hot to cold, look out below" doesn't really apply in Australia where the continental low pressure systems (rather than those emanating from the Southern Ocean) are not 'cold core lows' but 'surface heat lows' and troughs. See 'Height contours and thickness charts'. These altimeter indicated altitude variations should not be a concern to pilots of aircraft flying under the day visual flight rules and maintaining visual meteorological conditions, particularly so if en route area/local QNH baro-setting information is acquired and properly applied. What should be a particular concern is density altitude rather than true altitude. Calibrated altitude: is the altimeter indicated altitude, corrected for internal instrument error and static vent position error by means of reference data for that aircraft installation. Pressure altitude: is the altimeter reading when the baro-setting scale is set to 1013.2 hPa; usually termed pressure height in reference to an airfield reading. It is the ISA standard pressure setting. Standard pressure is also the standard factory setting for altitude encoding devices. All aircraft cruising in the Standard Pressure Region — above a transition altitude that (in Australia) commences at 10 000 feet — use the standard pressure setting, and the subsequent altimeter reading is normally referred to as flight level [FL]. True altitude: the calibrated/indicated altitude corrected for the outside air temperature conditions. However, there are still problems in the determination of the true height above the AHD, as demonstrated in the following paragraphs. True altitude as calculated in flight from an altimeter reading is of little value to recreational aircraft operating in VMC. GPS altitude: the global positioning system uses the WGS84 ellipsoid as its basis for GPS altitude, whereas the AHD (a 'geoid') is the basis for elevations on Australian navigation charts. The difference in elevation of a particular point on the Earth's surface — when measured against both the ellipsoid and a national geoid — can be quite considerable, as much as ±250 feet ; this is known as the geoid-ellipsoid separation. In Australia the degree of geoid-ellipsoid separation is quite unusual, in the south-west corner the AHD geoid is about 102 feet below the WGS84 ellipsoid while at the north-east corner it is 237 feet above it, so the value of the geoid-ellipsoid separation at all locations must be available to derive true altitude. See 'Geoid-ellipsoid separation and GPS altitude'. Density altitude: a calculation used to determine possible aircraft performance — see section 'High density altitude' below. This is the pressure altitude adjusted for variation from standard temperature, or the height in ISA having a density corresponding to the location density, then called density height. Declared density altitude: see 'Method 3: use the CASA declared density altitude charts' below. Pivotal altitude: is not associated with altimeter setting; it is a term used by the proponents of 'ground reference' manoeuvres such as 'eights on pylons'. It is a particular height above ground at which, from the pilot's viewpoint, the extended lateral axis line of an aircraft doing a 360° level turn (in nil wind conditions) would appear to be fixed to one ground point, and the aircraft's wingtip thus pivoting on that point. The pivotal altitude in nil wind conditions is easily calculated by squaring the TAS in knots and dividing by 11.3. So an aircraft circling at 80 knots would have a pivotal altitude around 550 feet, no matter what the bank angle. When an aircraft is turning at a height greater than the pivotal altitude, the wingtip appears to move backwards over the landscape. When an aircraft is turning at a height less than pivotal altitude (i.e. usually close to the ground) the wingtip appears to move forward over the landscape. For more information see 'pivotal altitude and reversal height'. Q-codes Note: the letters in the Q-code nomenclature have no literal significance; these are remnants of an extensive notation system from the days of wireless-telegraphy and particularly used in marine and aerial navigation/communication. There were some 200 three-letter Q-codes, each representing a sentence, a phrase or a question. For instance, QRM "I am being interfered with"!. Some 30 Q-codes are still used by amateur radio/morse code enthusiasts and the four below, plus QDM (the magnetic bearing to a station), still survive in aviation. For a full listing of Q-codes google 'all Q codes'. The following four codes relate to altimeter settings. QFE: the barometric pressure at the station location or aerodrome elevation datum point. If QFE is set on the altimeter pressure-setting scale while parked at an airfield, the instrument should read close to zero altitude — if the local pressure is close to the ISA standard for that elevation. However, the use of QFE is deprecated and anyway, if the airfield elevation is higher than perhaps 3000 feet, older/cheaper altimeters may not be provided with sufficient sub-scale range to set QFE. QFF: the mean sea-level [msl] pressure derived from the barometric pressure at the station location. This is derived by calculating the weight of an imaginary air column extending from the location to sea-level — assuming the temperature and relative humidity at the location are the long-term monthly mean, the temperature lapse rate is ISA, and the relative humidity lapse rate is zero. This is the method used by the Australian Bureau of Meteorology; QFF calculations differ among meteorological organisations. QFF is the location value plotted on surface synoptic charts and is closer to reality than QNH, though it is only indirectly used in aviation. QNH: the msl pressure derived from the barometric pressure at the station location by calculating the weight of an imaginary air column extending from the location to sea-level — assuming the temperature at the location is the ISA temperature for that elevation, the temperature lapse rate is ISA and the air is dry throughout the column. The Australian aviation regulations state that when an 'accurate' QNH is set on the pressure-setting scale at an airfield, the VFR altimeter indication should read within 100 feet of the published airfield elevation, or 110 feet if elevation exceeds 3300 feet; otherwise the altimeter should be considered unserviceable. However, due to the inherent inaccuracy possible in QNH, this may not be so. The difference between QFF and QNH when calculated on a hot day at a high airfield in Australia can be as much as 4 hPa, equivalent to about 120 feet. The advantage to aviation in using the less realistic QNH is that all aircraft altimeters in the area will be out by about the same amount, and thus maintain height interval separation. The local QNH at an airfield is normally derived from an actual pressure reading. But the area QNH used outside the airfield zone is a forecast value, valid for three hours, and may vary by up to 5 hPa from any local QNH in the same area. Either local QNH or area QNH may be set on the altimeter pressure-setting scale of all aircraft cruising in the Altimeter Setting Region, which (in Australia) extends from the surface to the Transition Altitude of 10 000 feet. The cruising levels within the Altimeter Setting Region are prefixed by 'A'; e.g. A065 = 6500 feet amsl. When there is no official Local QNH available at an airfield and the site elevation is known, the local QNH can be derived by setting the sub-scale (when the aircraft is on the ground at the location) so that the altimeter indicates the known airfield elevation. The use of local QNH is important when conducting operations at an airfield, as the circuit and approach pattern is based on determining height above ground level [agl]. Note that it is not mandatory for VFR aircraft to use the area QNH whilst enroute. You may substitute the current local QNH of any aerodrome within 100 nm of the aircraft or the local QNH at the departure airfield. See 'Acquiring weather and QNH information in-flight'. The purpose of the Transition Layer is to maintain a separation zone between the aircraft using QNH and those using the standard pressure setting. Cruising within the Transition Layer is not permitted. If Area QNH was 1030 hPa, there would be about 500 feet difference displayed between setting that value and setting standard pressure. The Transition Layer extends from the Transition Altitude to the Transition Level which, in Australia, is usually at FL110 but it may extend to FL125 — depending on Area QNH. More detail is available in 'Aeronautical Information Publication (AIP) Australia' section ENR 1.7; downloadable from Airservices Australia. QNE: common usage accepts QNE as the ISA Standard Pressure setting of 1013.2 hPa. However another definition of QNE is the 'altitude displayed on the altimeter at touchdown with 1013 set on the altimeter sub-scale' i.e. 'pressure height'. It is also referred to as the 'landing altimeter setting'. Within the latter meaning, the term is only likely to be used when an extremely low QNH is outside an aircraft's altimeter sub-scale range, and the pilot requests aerodrome QNE from air traffic services. In Australia, such extreme atmospheric conditions are only likely to occur near the core of a tropical depression/cyclone and as QNE is not listed in the ICAO "Procedures for Air Navigation Services", air traffic services would not provide QNE on request. However, QNE can be calculated by deducting the QNH from 1013, multiplying the result by 28 (the appropriate pressure lapse rate per hPa) and adding the airfield elevation. For example: QNH 960 hPa, airfield elevation 500 feet, pressure setting 1013. QNE = 1013 –960 = 53 × 28 = 1484 + 500 = 1984 feet (the reading at touchdown). 3.5.3 High density altitude: effect on take-off/landing performance High 'density altitude' conditions at an Australian airfield can provide a severely hazardous environment for any aircraft where the difference between power required and power available is small. This concerns most general aviation and all sport and recreational aircraft engaged in take-off or landing at that airfield. It is the density of the air that provides engine power, propeller performance and lift. What we are really doing when calculating density altitude is estimating the density of the air. In ISA conditions, at a density altitude of 6000 feet amsl, the air density will be about 1.0 kg/m³, about 20% less than sea-level standard density. The maximum lift possible to be generated will be reduced by 10% (lift = CL × ½rV² × S ) and the ground roll speed related to IAS/CAS prior to take-off will be greater; i.e. during take-off at msl in ISA sea-level conditions TAS = IAS/CAS, but in high density altitude conditions TAS is greater than IAS/CAS. Remember that V² in the lift equation refers to TAS not CAS. So, at 6000 feet density altitude the TAS at lift-off would be about 10% higher (see rule of thumb) than msl conditions thus the aircraft has to accelerate to a 10% higher ground roll speed before reaching lift-off IAS, and that is before taking into account the effect of the engine and propeller performance reductions on the aircraft's ground roll acceleration performance and its climb-out capability. The weight of the charge delivered to the cylinders, in a normally aspirated engine, will be only 80% of the standard sea-level value. Thus, only 80% of the engine's rated power can be supplied at the propeller shaft for take-off and climb-out, or for a go-around. The lower air density ( ½r in the ½rV² term of the lift equation) directly reduces the thrust performance of the fixed-pitch propeller by 10% in which case the thrust performance will be 90% of 80%, or about 72% of the rated sea-level performance. So both the time and the distance needed to acquire take-off lift — and to clear obstacles at the end of the strip — must be increased, the aircraft's rate of climb, and thus angle of climb, will be less than it is near sea level. There are many conditions that exist, or might exist, at high density altitude which, though they may be individually slight, all affect the airframe and engine performance adversely. For instance, attempting take-off with a combination of some of the following conditions may cause some difficulty; attempting take-off when most conditions exist may well be disastrous: at an elevated airfield with moderate to high surface temperature on a short, soft strip with unslashed, wet grass at maximum weight incorrect flap setting light and variable winds departing into rising terrain and a sinking air environment. The same conditions apply when landing; the TAS at Vref will be higher and the consequent ground roll will be longer. The thrust available for a go-around, in the event of an aborted approach, might be very much less than the rated msl thrust, which would probably preclude any late go-around. In addition, under high density altitude conditions, the mixture may be excessively over-rich. The recommendation for normally aspirated engines with cockpit mixture control is that the mixture should be leaned to maximum rpm before taxying, take-off or landing if the density altitude is 5000 feet or greater. Density altitude at a particular location can vary considerably from day to day, and also according to time of day. For instance, the table below shows a mid-afternoon and an early morning reading at Alice Springs, in central Australia, on different days. The airfield elevation is 1900 feet. QFE Temperature Air density Pressure altitude Density altitude 941 hPa 43 °C 1.037 2020 feet 5600 feet 957 hPa –2 °C 1.230 1580 feet –100 feet The isotherms plus colour in-fills on the following Australian Bureau of Meteorology map indicate the mid-afternoon surface screen temperatures on a late-spring day. Note that, except for the mountain area near the south-east coast, the surface temperatures greatly exceed the 15 °C ISA standard. 3.5.4 Calculating the dry air density altitude The density of dry air (r) varies according to ambient air pressure and ambient air temperature, this is reflected in the equation density = pressure divided by 2.87 times the temperature(K). Pressure (or pressure altitude) is readily obtained from the altimeter, and temperature can be obtained from various sources. Method 1: use the temperature differential Density altitude is roughly 120 feet greater than pressure altitude for each 1 °C that the temperature exceeds ISA for that level, and 120 feet less for each 1 °C that the outside air temperature is less than ISA. In the ISA table sea-level temperature is 15 °C and the ISA temperature lapse rate is 2 °C per 1000 feet. For example: Armidale, New South Wales, airport (elevation 3550 feet) on a warm summer day, temperature 30 °C. Altimeter, with 1013.2 standard sea-level pressure sub-scale setting, reads 3400 feet pressure height/altitude. So, ISA standard temperature for an elevation of 3550 feet = [15 –(3.55 x 2)] = 8 °C. The Armidale temperature then exceeds standard by 22 °C, thus adjustment to be added= 22 × 120 = 2640 feet Pressure altitude = 3400 feet Then the approximate density altitude = 2640 + 3400 = 6040 feet. Method 2: calculate using the air density equation The density of dry air at altitude can be calculated using the equation: r = P / (2.87 T), where: r = rho — the density of dry air [kg/m³] P = the pressure [hPa] 2.87 = the gas constant for dry air T = the air temperature in kelvin units [K]. Using the Armidale example, with the altimeter set so that altitude shows the elevation of 3550 feet, the pressure-setting sub-scale will display 895 hPa (i.e. QFE). The temperature is 303 K (30 °C + 273) thus density = 895 / (2.87 × 303) = 1.029 kg/m³. The height in ISA having a corresponding density is about 5850 feet. This gives a slightly more accurate calculation of density altitude than method 1. Method 3: use the CASA declared density altitude charts The ICAO International Standard Atmosphere model, used for flight instrument calibration, is based on average climatic conditions at 40° to 45° N latitudes and as such does not reflect conditions over much of Australia in all seasons, with the discrepancy peaking in summer. The Civil Aviation Safety Authority recognises this and publishes seasonal 'declared density altitude' charts with isopleths delineating regional values to be added to airfield elevation to give declared density altitude. The three seasonal charts (summer, winter and autumn/spring) are published as appendices to Civil Aviation Order 20.7.0. For example the summer chart shows regional values of 2000 feet in some south-east and 3600 feet in central areas. These regional values are to be used only if there are no other means of calculating current density altitude at the departure and destination airfields. Armidale is located at 30° S and 151° E between the 2800 and 3000 feet isopleths of the summer chart, so adding 2900 to the airfield elevation of 3550 feet gives a declared density altitude of 6450 feet. Method 4: use a density altitude computational chart First determine current pressure altitude with 1013 hPa standard pressure setting on the altimeter sub-scale, for example 3400 feet. Also determine outside air temperature, for example 30 °C. Draw a horizontal line from the 3400 feet position to the 30 °C vertical line. Determine the density altitude scale at which the line terminates, for this example density altitude = 6000 feet. Click the image for a larger scale printable computational chart. Method 5: use an E6-B type circular scale computer The plastic circular slide rule flight planning computers have a density altitude facility that just entails placing the pressure altitude opposite temperature and reading off density altitude from a third scale. Because the scales are close to the centre of the instrument they are small and difficult to set accurately, but the Jeppesen Model CR-3 computer indicates about 6000 feet density altitude for the Armidale example. It can be seen that the five methods listed provide much the same result — about 6000 feet density altitude — so use the method which is most convenient; but do the estimate and then calculate the effects on aircraft performance at estimated weight while also including the effects of runway conditions and the wind velocity. Method 6: use the information in the Pilot's Operating Handbook Charts in the Pilot's Operating Handbook or Flight Manual should provide density altitude plus aircraft performance and maximum weight figures from the input of pressure altitude and temperature. Method 7: use one of the density altitude calculators available on the internet Google "density altitude calculator". It's probably advisable to do a number of trial runs before choosing a particular DA calculator. Know the normal take-off distance required Before you can start to estimate the take-off distance required under high density altitude conditions, you must know the take-off distance required under standard ISA mean sea-level conditions. CAO 101-28, an airworthiness certification requirements for commercially supplied amateur-built kit ultralights, states in part: "The take-off distance shall be established and shall be the distance required to reach a screen height of 50 feet from a standing start, with ... short, dry grass surface ... the aeroplane reaching the screen height at a take-off safety speed not less than 1.2 Vs1 ... take-off charts ... shall schedule distances established in accordance with the provisions of this paragraph, factored by 1.15." Sea-level ISA and nil wind conditions are implied. CAO 101.55 has much the same wording, but specifies 1.3 Vs1 as the take-off safety speed. FAR Part 23 is similar. CAO 101-28 also requires that the landing distance stated will be that to come to a full stop from a screen height of 50 feet at the threshold, with the screen being crossed at 1.3 Vso and the same conditions as specified for the take-off distance. Refer to Vref. If buying an aircraft or kit, you should require that the standard take-off and landing distance chart information for the airframe/engine/propeller combination be supplied. Statements such as "Take-off ground roll 10 m to 40 m" have no value. You must insist, particularly with imported aircraft, that the distances should be stated clearly in one form only and for nil wind conditions "Take-off distance to clear 50 feet (15 m) screen" or "Landing distance over 50 feet (15 m) screen". You have to know without doubt, having done the necessary calculations, that you can clear obstacles at the end of the unslashed paddock on a hot, bumpy day without risk to you or your passenger, and that if it is necessary to abort a landing, the aircraft will have the ability to go-around safely. 3.5.5 Calculating and adding the effect of humidity on density altitude The ISA is based on dry air and though air density (mass per unit volume) is chiefly dependent on temperature and pressure, humidity — the presence of water vapour — does decrease the density of the air a little and so has a small effect on lift. Humidity also has a small adverse affect on engine performance as combustion performance is dependent on the oxygen intake during each engine cycle, and that amount of oxygen is dependent on the air density. The molecular mass ratio of water vapour to dry air is 0.62:1 and the water vapour molecule occupies about the same space as any oxygen or nitrogen molecule it displaces, so air density decreases as the relative humidity of the air increases, and this should be considered when calculating density altitude. *Note: see atmospheric moisture for more information. Effect on density altitude. The table below gives the density in grams per cubic metre of the water vapour at the saturation point (i.e. relative humidity (RH) = 100%), for air temperatures between zero and 45 °C. As can be seen, at 35 °C and 100% RH the water vapour density is 40 grams/m³ which is 3.5% of the dry air mass at that temperature. However as the mass ratio of water vapour to dry air is 0.62:1, the 40 grams of vapour in the moist air would displace 65 grams of dry air (13g of oxygen). The effect on air density is a net reduction of 25 grams or 2% and, below 5000 feet elevation, equivalent to a density altitude increase of about 750 feet. At 25 °C and 100% RH the water vapour density is 25 grams/m³ which would displace 40 grams of dry air (8g of oxygen) and the effect on air density is a net reduction of 15 grams or 1.2%; equivalent to a density altitude increase of about 500 feet. It can be seen that high humidity has an additional detrimental effect on aircraft take-off and landing performance under high density altitude conditions. Water vapour saturation partial pressure, density at sea level and effect on density altitude (values rounded) Air temperature Saturation p/pressure hPa Vapour density grams/m³ Dry air density grams/m³ Net density reduction grams/m³ Net density reduction % Density altitude increase 0 6 5 1290 3 0.2% +100 ft 10 °C 12 10 1250 6 0.5% +200 ft 15 °C 17 15 1225 9 0.8% +250 ft 20 °C 23 20 1200 12 1.0% +350 ft 25 °C 30 25 1180 15 1.3% +450 ft 30 °C 42 30 1160 19 1.6% +550 ft 35 °C 56 40 1150 25 2.2% 750 ft 40 °C 73 50 1130 31 2.7% +900 ft 45 °C 97 65 1110 40 3.6% +1200 ft The Australian Bureau of Meteorology publishes maps of the average monthly relative humidity observations and these might be used as a basis for estimation if you are unable to find the current RH at the location. If the dry-bulb and wet-bulb temperatures are known the following table will provide a reasonable estimate of the current relative humidity. Wet-bulb temperatures are always lower than dry-bulb temperatures. Calculation of relative humidity, for dry-bulb temperatures from 20°C to 45°C, knowing the difference between the dry-bulb and wet-bulb temperatures Difference -1° -2° -3° -4° -5° -6° -7° -8° -9° -10° Relative humidity 95% 90% 85% 80% 75% 70% 65% 60% 55% 50% A handy rule of thumb to allow for the effect of humidity — after completing the dry air density altitude calculation: for air temperatures below 30 °C add 50 feet to the dry air result for each 10% that the local relative humidity exceeds 10% e.g. air temperature 25 °C, RH 60%, altitude to be added is 5×50=250 feet for air temperatures 30 °C and above, add 100 feet to the calculated density altitude for each 10% RH that RH exceeds 20%; e.g. air temperature 35 °C, RH 65%, altitude to be added is 5×100=500 feet. The rule is based on the following calculations providing the increase in density altitude for air temperatures from 15 °C to 40 °C and relative humidities between 10% and 100%: 100% 90% 80% 70% 60% 50% 40% 30% 20% 10% 15 °C +260 ft +230 ft +210 ft +180 ft +160 ft +130 ft +100 ft - - - 20 °C +350 ft +320 ft +280 ft +250 ft +210 ft +180 ft +140 ft +100 ft - - 25 °C +450 ft +400 ft +350 ft +300 ft +260 ft +220 ft +180 ft +130 ft - - 30 °C +560 ft +500 ft +450 ft +390 ft +340 ft +280 ft +220 ft +170 ft +110 ft - 35 °C +740 ft +670 ft +600 ft +530 ft +450 ft +380 ft +300 ft +220 ft +150 ft - 40 °C +900 ft +810 ft +720 ft +630 ft +540 ft +450 ft +360 ft +270 ft +180 ft - For more information on take-off and climb performance in high density altitude conditions, see take-off considerations. 3.5.6 Physiological effects of altitude The tissues and organs of the human body need a constant and adequate supply of oxygen to function at maximum efficiency; insufficient oxygen in those tissue and organs is called hypoxia. There are many causes for the condition, but the one of most interest to sports and recreational aviators is the hypobaric form of hypoxia caused by continuing flight at an altitude where the partial pressure of the atmospheric oxygen is less than that required for proper functioning of the brain. The body utilises the oxygen partial pressure to pass it through the membrane of the lung alveoli into the bloodstream. (The 'stagnant' forms of hypoxia — greyout and blackout — caused by reduced blood flow to the eyes and brain at aircraft accelerations exceeding +3g to +4g is also, of course, of interest to aerobatic pilots. For a pilot of average fitness, greyout (dimness of vision) will start between +3.5g and +4.5g, reaching blackout (complete loss of vision) between +4g and +5.5g and g-induced loss of consciousness [GLOC] between +4.5g and +6g.) The application of perhaps –2g or –3g causes increased blood flow to the eyes, resulting in leakage from the blood vessels –redout. Prolonged application of high negative g may severely damage the optic nerves. Atmospheric oxygen partial pressure declines as altitude increases; see the atmospheric oxygen section in the Aviation Meteorology Guide. The table in that section shows the time a reasonably fit person will remain conscious at those altitudes without using supplemental oxygen. However, the effects of hypoxia commence at much lower altitudes, probably around 8000 feet for a fit person, less if unfit though much lower for a heavy smoker. These effects include a gradual deterioration in thinking, calculating and reacting; inability to make appropriate judgements; light headedness and a poor memory recall. Unfortunately, the afflicted person is usually quite unaware of the symptoms occurring and may enjoy a feeling of well-being even, perhaps, euphoria. For more information read the article 'Hypoxia' from Flight Safety Australia magazine. In Australia recreational aircraft may only be flown at or above 10 000 feet amsl if the pilot has applied to and received written permission for that flight from the Civil Aviation Safety Authority. The aircraft must be equipped with an operating Mode A/C or S transponder. Also the Australian Civil Aviation Order Part 20.4 paragraph 6 which applies to all Australian aircraft, requires that: "A flight crew member who is on flight deck duty in an unpressurised aircraft must be provided with, and continuously use, supplemental oxygen at all times during which the aircraft flies above 10 000 feet altitude." Note that an aircraft may not cruise within the transition layer and that layer could extend to FL125. Things that are handy to know Altimeter rules of thumb • For each 10 °C that the outside air temperature is warmer than ISA standard, increase the indicated altitude by 4% to give true altitude. Conversely, for each 10 °C cooler, decrease indicated altitude by 4% — 10/273 approximates to 4%; refer to Charles' law. • When flying from higher to lower pressure conditions, without altering QNH, the altimeter will — if below 10 000 feet — overread (indicate higher than actual altitude) by about 30 feet for each one hPa pressure change. • When flying from lower to higher pressure conditions, without altering QNH, the altimeter will — if below 10 000 feet — underread (indicate lower than actual altitude) by about 30 feet for each one hPa pressure change. • If the altimeter sub-scale setting is less than QNH the altimeter will overread. Conversely, if the setting is greater than QNH, the altimeter will underread. • Air density decreases by about 1% for each: — 10 hPa fall in pressure, or — 300 feet increase in height, or — 3 °C increase in temperature, from the msl standard. Stuff you don't need to know • There is a semi-diurnal atmospheric tide, similar to the oceanic tide, which is most apparent in the lower latitudes. The tide peaks at 1000 hrs and 2200 hrs local solar time, with the minima at 0400 hrs and 1600 hrs. At Cairns, 17° S latitude, the daily minima and maxima are 2 hPa either side of the mean pressure; e.g. 0400 hrs — 1014 hPa; 1000 hrs — 1018 hPa; 1600 hrs — 1014 hPa; 2200 hrs — 1018 hPa. The runway elevation at Cairns is 10 feet amsl, so that if you left a parked aircraft at 1600 hrs with the altimeter reading 10 feet, six hours later it would be reading 110 feet below mean sea-level. When making their regular pressure reading reports, weather observation stations adjust the reported QFF according to a 'time of day' table. • There is also a semi-diurnal gravity variation at the Earth's solid surface, also peaking at 1000 hrs and 2200 hrs. A movement of 50 cm from the low to high earth tide has been ascertained in central Australia. • Perhaps the highest surface pressure recorded is 1083.3 hPa at Agata, Siberia on 31 December 1968. Agata is 850 feet amsl. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com) Can not be reproduced without our express written consent

3.4.1 Variometer A variometer is a fast response vertical speed indicator usually scaled to match typical glider rates of climb and descent (+/-10 knots or +/- 5 metres per second). The variometer makes soaring possible by displaying the glider rate of climb to the pilot in near real time, enabling the pilot to manoeuvre the glider so as to remain in rising air. Variometers come in many types, some sense the airflow from a capacity bottle or chamber (as the outside static pressure increases or decreases due to altitude changes, air flows in or out of the chamber to equalise the pressure) either mechanically or electrically; others measure the air pressure directly using silicon pressure transducers and compute rate of climb electronically from the changes measured. All Borgelt variometers since 1982 use this last method. Audio signals which vary with rate of climb/descent are also possible when electronic sensing is used and this is a great help in keeping pilots looking outside the cockpit for other traffic, gliders or birds climbing better in nearby thermals and interesting and/or useful meteorological phenomena. 3.4.2 Vario averagers When circling in turbulent thermals – while our variometer will help us find the best lift – it is sometimes difficult to know how fast we are really climbing (or whether we really are climbing) as there may be sink in part of the circle and lift in others. In this case we can use an averager which is really just a slow response variometer. Most vario averagers integrate or average the variometer readings so that the running average rate of climb of the last circle or so is shown. This requires averaging over 20 to 30 seconds. The averager display may be "on demand" or continuously displayed in digital form. 3.4.3 Total energy variometer The basic variometer described above suffers from the effect that as the glider changes airspeed in response to pilot inputs, large transient rates of climb and descent are induced until the airspeed is stabilised at a new value. These may easily exceed and swamp the rates of climb due to airmass rising and sinking (a 30 degree pullup from 100 knots gives an initial rate of climb of 50 knots, causing the +/- 10 knot scale variometer to peg uselessly at the top of its scale). If, instead of sensing the outside air pressure or static pressure with the variometer, we connect the variometer to a venturi of the correct dimensions we find that as long as the airspeed is constant the pressure in the venturi decreases and increases as in the basic variometer case. If the airspeed decreases due to a pullup the suction produced by the venturi will decrease and will compensate for the reduced static pressure from the climb resulting in no net change of pressure and hence no change in variometer reading. The suction produced by this venturi (pressure below static pressure) is the same as the pressure increase above static pressure measured by a pitot tube at the same airspeed. For the last 25 years the most common and best "venturi" in fact doesn't look like a venturi at all. A 6 mm tube usually extends from the fin leading edge and is bent up or down 70° or so, so that the last 80 mm of the tube is at 20° forward inclination to the airflow. The end of the tube is sealed and the end is cut off square to the tube (NOT parallel to the airflow) and two small holes are drilled in the rear half of the tube as pressure ports. The suction of the device depends on the distance of the holes from the end of the tube. This design is relatively insensitive to yaw (sideslip) and pitch and unlike a real venturi doesn't provide a home for spiders and insects. It is also easy to keep clean (required for correct functioning). As it was invented by Frank Irving it is known as an Irving tube. The generic term for all venturis, probes etc is Total Energy Probe or TE Probe. The total energy referred to is kinetic plus gravitational potential energy. A total energy variometer as described above can be further improved. As described the vario will, in still air, show the glider sink rate at the speed being flown. Let us take a good modern glider such as an 18 meter racer. The unballasted polar curve for this glider will show a minimum sink rate of around 1 knot at around 45 knots IAS, a best L/D of about 50 at around 55 knots and sink rates of about 2 knots at 75 knots and 4 knots at 100 knots. The sink rate through the airmass at typical thermalling airspeed and bank angle will be about 1.6 knots. Now suppose you are cruising between thermals and you encounter rising and sinking air and you vary the airspeed according to Macready 'speed to fly' theory between 60 knots and 110 knots as you encounter rising and sinking air. The vario is TE compensated so changes in airspeed don't cause large transient indications on the vario. However at 110 knots the glider might be sinking through the airmass at 5 knots and at 60 knots at just over 1 knot. These sink rate changes can mask small vertical speed changes in the airmass and make it difficult to pick the best path through the air (which is the path with the most and fastest rising air and the least and slowest sinking air). If encountering a thermal at high speed you may even reject the thermal that is really acceptable. A thermal rising relative to the ground at 8 knots will show as 3 knots up on the vario if you are sinking at 5 knots at 110 knots IAS. In fact after you reduce speed and turn in this air you would climb at 8 – 1.6 = 6.4 knots so it is difficult to use the vario indication to decide whether to accept or reject the thermal being flown through. 3.4.4 Netto variometer A Netto or airmass variometer adds to the raw TE vario reading an upward deflection to counteract the sink due to the glider polar at that airspeed. Now, in still air, the vario will read zero at any airspeed if the polar we have assumed is correct. Even if it isn't the differences are likely to be small fractions of a knot and it is much easier to use the vario to pick the best path through the air. There is one disadvantage and that is if we fly through our thermal rising at 8 knots relative to the ground we see 8 knots on the Netto vario regardless of the airspeed we are flying at. When we slow to circle we climb at 8 – 1.6 = 6.4 knots. So we mentally must subtract our circling sink rate of 1.6 knots to see what rate of climb we will get if we circle now. This is much easier than with our raw TE vario but still adds to workload (for most gliders 2 knots is in fact close enough for practical purposes). 3.4.5 Relative Netto variometer A further refinement of the Netto vario and is sometimes called just relative or super Netto. If we superimpose a downward deflection of 1.6 knots on our Netto variometer we can see that at any airspeed the variometer will show the rate of climb we will get if we circle. This now makes it very easy to use the vario to decide if the thermal is good enough (it isn't the only criterion – you may not have flown through the center – you may already be low etc) but at least the vario is easy to interpret. For picking the best path through the air the relative vario is almost as good as the Netto. Any time that the vario is heading in the upward direction the air is getting better, downwards worse. To get the best of both the Netto and relative vario just mark the 2 knots down position on the outside of the scale with a sliver of white tape. This serves as the "still air" reference point. Any time the vario is above this point the airmass is rising. Do we care about this? Surely we just want the best air? Sometimes it is important in interpreting the meteorology to know if the air is rising or sinking. Also our estimate of how well the glider will go on final glide depends on knowing the absolute rise or fall of the air we fly through. We may be flying in the best available air but if that airmass is sinking at only a fraction of a knot on average we must allow extra altitude for final glide. 3.4.6 Speed command variometer Earlier in the article I mentioned Macready 'speed to fly' theory. Depending on the anticipated strength of the next thermal and the glider polar and the air you are currently flying through there is an optimum indicated air speed to be flying at. A table of speeds is one way to do this, a moveable scale (Macready ring) around the vario is another. If we have the electronics to measure our airspeed to provide a Netto or relative vario we can also use a little more processing to provide a zero reader for the optimum speed to fly. Pointer above zero - pull up, reduce speed. Below - push and gain speed. This is known as a speed command variometer or speed command. Audio signals for "fly faster" and "fly slower" may also be generated. This combines the TE variometer reading with a downward offset controlled by the MaCready or STF (speed to fly) selector (may be a rotary knob or controlled by accessing a menu on screen) and an upward deflection that gets larger as airspeed is increased. If done correctly the result is that you fly at the optimum speed at all times. The speed changes should be made gently so the zero reader indications are usually filtered to slow down the commanded changes. 3.4.7 Effects of altitude It should be noted that the pressure change for any given altitude increment reduces with altitude so variometer calibrations require correction for the effects of altitude. The TE probe will automatically compensate for the effects of altitude. Because the air is less dense at altitude which causes the glider polar to change, Netto, relative and speed command indications also require correction for altitude. The glider polar also changes with the weight of the glider and with the contamination of the airfoil by insects so both the glider weight and the degradation of the polar due to bugs must be accounted for in the Netto, relative and speed command indications. The processing in the B40 Vario and B50 Super Vario takes all the relevant effects into account. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.3.1 The atmospheric pressure gradient The random molecular activity or internal kinetic energy within a parcel of air is known as the static pressure and is proportional to the absolute temperature. Static pressure exerts a force on an object (for example, an aircraft wing) at right angles to all the exposed non-porous surfaces, and is measured in newtons per square metre [pascals] of surface. In Australia air pressure is reported as hectopascals [hPa] for meteorological purposes; one hectopascal equals 100 N/m² [or one millibar]. Atmospheric pressure reflects the average density (i.e. mass per cubic metre), and thus the weight, of the column of air above a given level. So, the pressure at a point on the Earth's surface must be greater than the pressure at any height above it, in that column. An increase in surface pressure denotes an increase in mass — not thickness — of the blanket of air above the surface location. Similarly, a decrease in surface pressure denotes a decrease in the mass of air. The air throughout the column is compressed by the weight of the atmosphere above it, thus the density of a column of air is greatest at the surface and decreases with increasing altitude. However, a warmer air column will be thicker — i.e. extend further upwards — than a cooler air column with the same surface pressure. Thus a particular pressure level will be at a higher elevation in the warmer column. This means that the level in the atmosphere at which any particular pressure occurs is also dependent on the temperature (or thickness) of the air column. Meteorological offices produce 'height contour and thickness charts' to determine the locations of upper level troughs and ridges. 3.3.2 Atmospheric density The average density of dry air in mid-latitude, temperate climates, is about 1.225 kg/m³ at mean sea-level; the density decreases with increasing altitude. There are several gas laws that relate the temperature, pressure, density and volume of air. The equation most pertinent to aeronautical needs is the equation of state: r = P/RT where: r* = air density kg/m³ P = static air pressure in hectopascals [hPa] R = the specific gas constant for dry air = 2.87 T = air temperature in kelvins [K] = °C + 273 *r is the Greek letter rho, pronounced 'row', as in 'row your boat'. By restating the equation of state as: P = 2.87rT , it can be seen that if density remains constant; pressure increases if temperature increases. We can calculate the ISA standard sea-level air density, knowing that standard sea-level pressure = 1013 hPa and temperature = 15 °C or 288 K i.e. Air density = 1013 / (2.87 × 288) = 1.225 kg/m³ = ro* If the air temperature happened to be 30 °C or 303 K at the same pressure, then density = 1013 / (2.87 × 303) = 1.165 kg/m³, or a 5% reduction. *ro — the symbol for the standard sea level density is pronounced 'rho zero', or 'rho nought' if you prefer. 3.3.3 The ICAO International Standard Atmosphere The International Civil Aviation Organisation's [ICAO] International Standard Atmosphere [ISA] provides a fixed standard atmospheric model that is used for many purposes, among which are the uniform assessment of aircraft performance and the calibration of some aircraft instruments. The model is based on average climatic conditions in 40–45° North latitudes, but contains the following assumptions: dry air (no water vapour present) is assumed throughout the atmosphere so the effects of humidity on air density are ignored the mean sea-level [msl] pressure = 1013.25 hPa the msl temperature = 15 °C [288 K] the tropopause is at 36 090 feet [11 km] and the pressure at the tropopause = 226.3 hPa the temperature lapse rate to 36 090 feet = 6.5 °C per km, or very close to 2 °C per 1000 feet the temperature between 36 090 and 65 600 feet [20 km] remains constant at −56.5 °C. The table below shows a few values derived from the ISA. Those pressure levels noted with a flight level designator (FL) are standard pressure levels, rounded to the nearest increment of 500 feet, used for aircraft operating above the altitude transition layer and also for aviation weather purposes, particularly thickness charts. hPa °C kg/m³ feet Pressure Flight level Temperature Air density Altitude 1013 15.0 1.225 msl 1000 14.3 1.212 364 977 13.0 1.190 1000 950 11.5 1.163 1773 942 11.0 1.155 2000 908 9.0 1.121 3000 900 8.6 1.113 3243 875 7.0 1.088 4000 850 A050 5.5 1.063 4781 843 5.1 1.056 5000 812 3.1 1.024 6000 800 2.3 1.012 6394 782 1.1 0.993 7000 753 – 0.9 0.963 8000 750 –1.0 0.960 8091 724 – 2.8 0.933 9000 700 A100 – 4.6 0.908 9882 696 – 4.8 0.905 10 000 650 – 8.3 0.855 11 780 600 FL140 –12.3 0.802 13 801 550 –16.6 0.747 15 962 500 FL185 – 21.2 0.692 18 289 450 – 26.2 0.635 20 812 400 FL235 – 31.7 0.577 23 574 350 – 37.7 0.518 26 631 300 FL300 – 44.5 0.457 30 065 250 FL340 – 52.3 0.395 33 999 200 FL385 – 56.5 0.322 38 662 150 FL445 – 56.5 0.241 44 647 100 – 56.5 0.161 53 083 It can be seen that (from sea level to 10 000 feet) air density decreases by about 33 grams/m³ per 1000 feet. Also, not immediately apparent from the ISA table, is that the pressure lapse rate (the rate of change with height) is exponential. It starts at about one hPa drop per 28 feet height increase, then slowing to 31 feet per hPa at 6000 feet (averaging 30 feet height increase per hPa drop up to 6500 feet), 36 feet per hPa at 10 000 feet, 50 feet at 20 000 feet and so on. However, the following provides a useful rule of thumb: Rule of thumb #1 "For operations below 10 000 feet, an altitude increase (decrease) of 30 feet can be assumed for each one hPa pressure decrease (increase) and for an estimate of air density multiply the altimeter indicated altitude in 1000's of feet by 30 to find the value to deduct from 1.225 kg, e.g. at 6500 feet; 6.5 × 30 = 195 and 1.225 kg minus 195 grams = 1.03 kg." But bear in mind that the ISA model is unlikely to reflect current atmospheric conditions at a particular location, see high density altitude. 3.3.4 Bernoulli's principle and the continuity equation Bernoulli's principle Daniel Bernoulli (1700-1782) was a Swiss mathematician who propounded the principle that for a given parcel of freely flowing fluid, the sum of gravitational potential energy, kinetic energy and static pressure energy always remains constant. For our aerodynamic purposes we can ignore the gravitational potential energy. Dynamic pressure = ½rv² and kinetic energy = ½mv² where m = mass. Air density is mass per unit volume; i.e. kg/m³ so dynamic pressure is the kinetic energy per unit volume. Static pressure is internal kinetic energy per unit volume, or pressure potential energy. So, for our purposes (in a parcel of freely flowing air), Bernoulli's principle can be reduced to: ½rv² [dynamic pressure] + P [static pressure] = constant The statement doesn't take into account viscosity, heat transfer or compressibility effects, but for operations below 10 000 feet and airflow velocities below 250 knots, compressibility effects can be ignored — thus no change in flow density [r] is assumed. The statement then indicates that, in a free stream flow, if speed [v] increases static pressure [P] must decrease to maintain constant mechanical energy per unit volume; and the converse — if speed decreases, static pressure must increase. Or, turning it around, a free stream airflow will accelerate in a favourable pressure gradient and decelerate in an adverse pressure gradient. Bernoulli's principle doesn't apply in boundary layer flow because the viscosity effects introduce loss of mechanical and thermal energy (transferred to the aircraft skin) due to the skin friction. (Incidently Daniel Bernoulli's father Johann [born 1667] was the mathematician who first adopted the symbol 'g' for the acceleration due to gravity.) Stagnation pressure Another aspect of Bernoulli's principle is that the constant is the stagnation pressure — the pressure energy needed to halt the airflow — thus it can be written ½rv² + P = stagnation pressure. The stagnation pressure is the highest pressure in the system. This application of the principle is apparent in the air speed indicator, as demonstrated below. Stagnation pressure is the basis of the ram-air parachute wing used in sport parachutes, paragliders and powered parachutes, see 'The ram-air parachute wing'. Also aircraft fuel-tank vents face into the airstream and rely on stagnation pressure to prevent inflight siphoning of fuel from the tank. The continuity equation There is another principle of aerodynamic interest to us — the fluid flow continuity equation — which states that, in a steadily moving airstream, the product of density, cross sectional area [s] and speed must always be a constant: r × s × v = constant If there is no change in density within the flow (which is the norm in the airspeed range of light aircraft; see compressibility effects) then we can state that: s × v = constant Thus, if air flows into a smaller cross-sectional area speed must increase to maintain the constant. Bernoulli's principle states that if speed increases, static pressure must decrease; so the velocity of a constricted airstream increases and its static pressure decreases. Both the above principles are related to the conservation laws; Bernoulli's principle to the conservation of energy, and the continuity equation to the conservation of mass. We will examine these properties of air further in the 'Aerofoils and wings' module. The venturi effect — used in carburettors, the total energy variometer and the airframe-mounted venturi that provides suction for some flight instruments — is an application of the principles stated above. 3.3.5 Measuring airspeed The dynamic pressure of the airflow, in N/m², is represented by the expression ½rV², where: r is the ambient density of the air [kg/m³ ] ( ro is the symbol for the standard sea-level density of 1.225 kg/m³ ) V² is the aircraft (or free airstream) speed [m/s²] and we can deduce that the apparent speed of the airstream is related to air density and dynamic pressure. *Note: a lower case v is the symbol for speed in physics while an upper case V is generally the symbol for the free stream speed in aerodynamics, which is why I have used the lower case v in section 2.4 but an upper case V in this section and most of this flight theory guide.) We can measure the dynamic pressure with a simple mechanical pressure gauge. Imagine a 6 mm internal diameter aluminium tube positioned under the wing of a moving aircraft, outside the slipstream, so that the open end points forward into undisturbed airflow and the other end of the tube terminates within a spring-loaded, flexible capsule — similar to that in an aneroid barometer — thus the capsule stops the airflow within the tube. The back pressure, applied by the capsule to stop the airflow, must be equal to the stagnation pressure. The capsule is contained within a casing which, in turn, is connected to a static vent that supplies the casing with the ambient atmospheric pressure; or, in a lower-quality system, the casing may just be open to the atmospheric pressure within the fuselage. So, if we have stagnation (or impact pressure or ram-air pressure) — which is dynamic pressure plus static pressure — within the capsule and static pressure surrounding it, the capsule will expand or contract to reflect the changes in dynamic pressure at the mouth of the tube. (The system is a 'pitot tube' devised by Henri Pitot (1695-1771). During World War I, the airspeed indicating instruments themselves were called 'pitots'.) The capsule movement is mechanically or electrically linked to rotate a pointer on a dial. Although the dial is calibrated and marked to indicate airspeed in knots or mph rather than hPa, it is still basically a simple pressure gauge and an imperfect airspeed gauge. Because the instrument cannot determine the density component of the dynamic pressure, the calibration assumes a constant air density of 1.225 kg/m³. This cockpit instrument is then an airspeed indicator [ASI] and it displays the indicated airspeed [IAS], based on ISA conditions. The indicated airspeed is not the actual aircraft speed through the air, 'V' in the equations. A bit confusing — but brace yourself, for it gets worse! We can calculate the dynamic pressure for the Jabiru using the scenario in section 1.4 for calculating CL; i.e. cruising at 6500 feet, true airspeed 97 knots or 50 m/s, and air density 1.0 kg/m³. The ISA atmospheric pressure at 6500 feet is about 800 hPa. • static pressure = 800 hPa • dynamic pressure = ½rV² = ½ × 1.0 × 50 × 50 = 1250 N/m² = 12.5 hPa Note that the dynamic pressure at 1250 N/m², or 12.5 hPa, is less than 2% of the static pressure, but applying that dynamic pressure over the 8 m² of wing area and the lift coefficient of 0.4; i.e. 1250 × 8 × 0.4, still gives the lift force of 4000 newtons that we calculated in the 'Lift' section. The airspeed 'V' in the equations is the true airspeed [TAS] — the free stream speed or the air distance flown over time. We know that the ASI dial is calibrated assuming a fixed air density of 1.225 kg/m³ [ro ], so a perfect ASI will only indicate the real airspeed (the true airspeed) when the actual environment density is 1.225 kg/m³; that could only occur when the aircraft is operating at low altitude. What will be the IAS (Vis) of our example in the preceding box? IAS = Vis = V / √(ro / r) = 97 divided by the square root of 1.225 divided by 1.0 = 97 / 1.1 = 88 knots From this we can deduce that a perfect ASI will generally underread. The IAS will always be less than the TAS, except in very cold conditions at low altitude where the air density may be greater than 1.225 kg/m³. For instance, using the equation of state above, if temperature was –3 °C and pressure was 1030 hPa, the density would be 1.33 kg/m³. Density is about 1% less than (greater than) the ISA value for each 3 °C that the temperature is above (below) the ISA value. 3.3.6 Indicated and calibrated airspeed The ASI as an angle of attack indicator So, you might ask, what's the point of an ASI that really is indicating just dynamic pressure, and is unlikely to indicate your real airspeed — air distance flown over time — accurately? Well, admittedly it does mean a little more calculation to be done in navigation, but there are very significant advantages with an instrument that displays IAS rather than TAS. This will be covered in the 'Aerofoils and wings' module. Generally, for all angles of attack in unaccelerated flight at a particular weight, there is a corresponding IAS; though the relationship between aoa and IAS does get a bit fuzzy near CLmax. So, in the absence of an angle of attack instrument, the ASI can generally* be regarded as an indication of angle of attack if the lift being produced matches aircraft mass. Also, all the performance parameters (the 'numbers') for an aircraft — best rate of climb, best angle of climb, best glide angle, etc. — require it to be flown at a particular aoa for that weight, and thus a particular IAS. Or more accurately, a particular calibrated airspeed [CAS] and that particular CAS does not change with altitude (as TAS does), but changes only with weight. *The reason why CAS does not always correlate to aoa in level flight is that when inertia and random displacement forces — atmospheric turbulence — come into play, aoa may change momentarily without a noticeable change in CAS. (Note: there are means of audibly conveying the angle of attack to the pilot. The simplest is a stall warning horn operated by a basic two-position vane switch incorporated in the leading edge of the wing and switched on by the airstream pressing the vane up if the aoa comes close to the stalling aoa. There are other airstream direction detectors (ADDs) that provide a range of warning tones in the pilot's headset.) Calibration of the aircraft's airspeed indicator system An ASI is an imperfect mechanical instrument which is subject to instrument errors; the poorer the quality of the instrument the greater the instrument error; the permissable limit for a certified ASI is ±one knot. The associated pitot/static and tubing system is also prone to pressure sensing errors due to the positioning (and design) of the the static air vent —and possibly the pitot head — relative to the airstream. Relative airflows change as aoa changes and when slipping or yawing. It is difficult to find a site on the fuselage for a static vent where the static pressure reflects the ambient static pressure. So after construction is complete the aircraft's airspeed indicating system must be 'measured' to determine the rectifications to be made to the airspeed reading indicated on the ASI. CAS is the airspeed after you have applied corrections to the IAS for those instrument and position errors occurring at that aoa in that particular aircraft. The measured corrections should be stated on a card placed near the ASI. You should also be aware that position errors may be quite significant, possibly under-reading by 10 knots or so — particularly at high aoa or when the pilot is maintaining a significant sideslipping manoeuvre. CAS may also be known as the rectified airspeed [RAS]. Below is an airspeed correction card for a particular aircraft in balanced level flight; i.e. not slipping or skidding. The normal cruise speed for this aircraft is around 95 knots. In this particular installation the ASI significantly underreads at low speeds and overreads at high speeds. IAS knots 42 52 61 69 73 87 96 104 113 122 130 CAS knots 49 57 64 71 73 86 94 102 110 117 125 Converting CAS to TAS TAS = CAS × √(ro / r) Using our example, from section 2.5, of the Jabiru cruising at 88 knots IAS at 6500 feet, where the air density is 1.0 kg/m²: TAS= 88 × √(1.225 / 1.0) = 88 × 1.107 = 97 knots We need TAS for navigation and as the density lapse rate just about follows a straight line below 10 000 feet – there is a simple mental calculation to determine TAS from CAS: Rule of thumb #2 "To convert CAS to TAS, multiply the (density) altitude, in 1000s of feet, by a factor of 1.5 to get the percentage increase to apply." (For calculation of density altitude see 'High density altitude: effect on take-off/landing performance'.) e.g. CAS = 88 knots at 6500 feet = 6.5 x 1.5 = 10% = 97 knots. The multiplication factor increases slightly with increasing altitude, reaching 2 at 30 000 feet. The airspeed indicator You will note the green and white peripheral arcs, and other colour marks, on the face of this instrument. These are standard markings, some of which should appear on the face of every light aircraft ASI, as they display the speed constraints applicable to the aircraft operations. The white arc indicates the flaps operating range starting, at the lower end, from the indicated airspeed, Vso [55 knots], at which the aircraft will stall in the landing configuration with flaps fully extended, and with the throttle closed. The top end of the white arc indicates the maximum speed, Vfe [108 knots], at which the aircraft's flaps can be extended, or remain extended, without causing strain. The bottom end of the green arc indicates the stalling speed of the aircraft, Vs1 [62 knots], with flaps (and landing gear if applicable) up, throttle closed and 1g load factor. The top end of the green arc indicates the maximum structural cruise speed, Vno [173 knots]. The green arc indicates the designed range of speeds for normal operations. The yellow arc indicates a speed range in which the aircraft may be flown, but with caution and only in smooth atmospheric conditions. The red line at the top end of the yellow arc indicates the speed, Vne, that should never be exceeded because of risk of structural damage. The other red mark at 70 knots, and the blue mark at 88 knots, are of no interest for single-engine aircraft; these markings only appear on a twin-engine aircraft ASI and relate to operations with one engine shut down. A properly functioning ASI responds rapidly to pressure changes because there is no instrument lag. A slow response attributed to instrument lag is most likely only due to the inertia of the aircraft — when attitude in pitch is changed, an aircraft takes a little time to accelerate/decelerate to the appropriate airspeed. Airspeed summary True airspeed [TAS] = V in the dynamic pressure equation and other expressions = air distance flown over time. Indicated airspeed [IAS] = Vis = airspeed displayed on the cockpit airspeed indicator [ASI] — based on a fixed air density ( ro ) of 1.225 kg/m³. The ASI only indicates true airspeed when ambient atmospheric density is actually 1.225 kg/m³ and the system error corrections are applied. Calibrated airspeed [CAS] = IAS adjusted (mentally from an airspeed correction card) for known system errors occurring within the normal speed range. Electronic ASI Electronic flight instrument systems [EFIS] use solid state electronic componentry as sensors plus software to display flight data on a single screen. In such systems, the static and dynamic pressures are fed to pressure transducers which sense and convert pressures to voltages that the electronic circuitry converts to an airspeed display. See the liquid crystal primary flight display of the Dynon D10A light aircraft EFIS. The EFIS has an outside air temperature probe and, with static pressure, the software can calculate air density and thus display TAS when needed. Electronic ASIs are also available as single panel instruments or possibly combined with an altimeter function. The electronic systems are still subject to much the same errors as a mechanical system, and the IAS has to be corrected for CAS unless there is a means for incorporating some form of compensating table into the software. 3.3.7 Measuring rate of ascent/descent The vertical speed indicator In flight it is important for a pilot to know the rate at which the aircraft may be ascending or descending. A simple vertical speed indicator [VSI] is a pressure gauge that measures the rate of pressure change as an aircraft is ascending or descending. The instrument display is usually calibrated in feet per minute but it may be in metres per second. There are two pressure inputs, both from the static vent system — one to each side of a flexible diaphragm or capsule. On the open side there is a normal input that reflects the static pressure change as it occurs. On the closed side the input/output is a fine capillary tube that slows the equalising pressure change — and also the response time of the instrument. The resultant deflection of the diaphragm is magnified via a geared mechanical linkage to a dial pointer, which indicates whether the aircraft is maintaining altitude (in which case, the pressure on both sides of the diaphragm is equal), climbing or descending, and the rate, usually graduated in feet [×100] of altitude per minute. Some form of vibration damping and thermal change compensation is included within the VSI and ASI instruments. The pressure change has two components, the most significant component is that brought about by the aircraft's rate of height change, as in climbing or descending. The other part is any vertical movement of that part of the air mass in which the aircraft is operating — rising air or sinking air. When an aircraft is climbing, rising air adds to the rate of change, sinking air reduces it. When descending, rising air decreases the rate of change, sinking air adds to it. The variometer In soaring flight, paraglider, hang glider, sailplane and power-assisted sailplane pilots are totally reliant on finding sources of atmospheric uplift to gain the gravitational potential energy that enables the aircraft to stay airborne for sufficient time to complete the flight plan. A variometer (usually abbreviated to vario) is a specialised vertical speed indicator that enables a pilot to derive the vertical speed of the parcel of air in which the aircraft is soaring. For more information on varios and their uses see the article 'Basic sailplane instruments. The article only refers to fixed installation varios but very light-weight hand-held varios are available for hang glider/paraglider pilots. 3.3.8 Stalling airspeeds The normal 1g stall One of the first questions a pilot might consider, when converting to a new aircraft type, is "What's the stall speed?" The reason for considering this is that usually, but not always, the approach speed chosen for landing is 1.3 to 1.5 times Vso — the minimum steady flight speed in the landing configuration, below which speed the aircraft will stall or at which speed the aircraft will stall if any manoeuvring is attempted. In aerodynamic terms, the 'stall' is the sudden widespread separation of the boundary layer from the upper wing surface that occurs when the wing exceeds a particular angle of attack. For light aircraft without high-lift devices, this is usually around 15–16° although minimum aircraft with single-surface fabric wings may have a stall aoa 2–3° lower. This critical angle of attack has no relationship with either the aircraft attitude relative to the horizon or the airspeed — it can readily be reached in a high-speed dive. But it does have a direct relationship with elevator position and thus the control column position. The separation of the boundary layer starts at the wing trailing edge, generally near the wing root for approximately rectangular wings (and particularly for wings with 'washout'), spreading forward and outward over the upper surface until there is a significant detachment of boundary layer flow over the upper surface. There is probably little change to the under-surface boundary layer flow. Between the two remnant boundary or shear layers, a thick turbulent wake will attach to the wing and be dragged along by the aircraft. The reaction to the acceleration and energising of that wake is a sudden deceleration of the aircraft accompanied by a large increase in the nose-down pitching moment plus some loss of lift. The initial wake turbulence ('burbles') near the wing root may initiate unsteady flow over the tailplane, shaking the tailplane and thus providing a few warning buffets felt in the airframe or smaller 'nibbles' felt in the controls — aerodynamic warnings of an impending stall. There also may be 'oil-canning' noises from pressure changes on metal-skinned fuselages and wings as the thin metal flexes in response to pressure changes. On the other hand, there may be no pre-stall warning whatsoever. Some aircraft exhibit undesirable characteristics even before boundary layer separation occurs; for example, the aircraft starts sinking excessively with increasing induced drag or wing rocking occurs. So, from a pilot's point of view, a stall is "the point following deceleration at which the pilot ceases to have full control over the aeroplane"; which adds the concept of the defined stall speed being a minimum controllable steady flight speed at which no undesirable characteristics are exhibited. The next comments are specifically aimed at stalls induced when: • flying straight and level at slower speeds • in a low speed descent — such as the approach to landing • in a climb — such as the initial climb after take-off • in a go-around following an aborted landing approach. The last two circumstances are sometimes referred to as full-power stalls or 'departure stalls'. In non-turbulent atmospheric conditions, and if the aircraft is in balance, all of the circumstances above can only induce a stall if the control column position is placed in, trimmed into or allowed to move into, the last half of its rearward travel. Many aircraft are designed so that the control column must be at or near the limit of its rearward travel to reach the stalling aoa. (The rearward travel range commences from the neutral position, as does the forward travel range.) Because of the airflow turbulence and increasing induced drag as the aoa is increasing, total drag increases and the aircraft slows as it approaches CL max. The rapid reduction in airspeed after passing the critical aoa means the wing is now unable to provide sufficient lift to totally balance weight and, in a normal stall, the aircraft starts to sink. The (possibly pronounced) nose-down pitch will occur even though the control column is near its rearward travel limit. However, some aircraft may not assume that nose-down attitude but just sink (mush down) at quite a high rate and at an extreme angle of attack. Because of the nose-up attitude, the high rate of descent may not be apparent unless the aircraft is close to the ground. The aircraft is instantly recovered from the stall by smoothly reducing the aoa so that it is below the critical aoa; i.e. easing the control column forward and generally no further than the neutral position. If one wing stalls before the other, that wing will drop. In this case, the control column must be firmly moved sufficiently forward to unstall the dropping wing, the wings levelled with aileron then sufficient rudder applied to stop further yaw. An increase in speed is also needed (by increasing power or holding a lower nose attitude for several seconds) so that a safe flight speed is achieved quickly without wasting much altitude and the aircraft is returned to the intended flight path. See standard recovery procedure for all stall types. If the control column movement for stall recovery is both excessive and abrupt, the result could be an aoa movement below the zero lift aoa — in which case there will be a reversed lift force on the wings, which hinders recovery. Weight-shift controlled trikes do not react well to negative g; if this is excessive, the wing spars may buckle at an outer position. Many aircraft are designed so that the nose will drop at the stall, but the aircraft will self-recover (i.e. without pilot intervention) in a stable descent or with some oscillations which, if the control column is still held back, will result in another stall. Some aircraft may be designed so that the wing is usually not able to reach the stalling angle, but the aircraft will enter a semi-stable mushing descent — which sounds fine but can be disastrous if the pilot doesn't notice when close to the surface. A normal stall occurs when the load factor is close to normal; i.e. near 1g. The cg position will also affect the manner of stall. If the cg is at the extreme forward limit, some aircraft may not fully stall — just mush down. If it is too far aft, the stall aoa can be reached with a much smaller rearward movement of the control column. Another factor affecting the manner of stall is the use of power. Generally, when flying slowly, the longitudinal axis of the aircraft is pitched up relative to the flight path. Consequently the thrust vector will include a vertical component — a lifting force — and the amount of lifting force provided depends on the amount of thrust. Also, for aircraft with the propeller mounted in front of the wings, the energy in the slipstream tube in slow flight increases the velocity of the airstream over part of the wing (depending also on the mounting of the wing in relation to the thrust centre line) and reduces the aoa of that part. Thus the completely stalled wing may occur at a lower speed, depending on the amount of power in use. When it occurs, the stall will be much more pronounced, possibly with a fast-acting wing drop. There are other complications because the slipstream also affects parasite drag and induced drag. Many pilots, in suitable aircraft and atmospheric conditions, prefer to land by approaching at 1.3 to 1.5 times normal stall speed — Vso — and, after flaring with the throttle closed, holding the aircraft just above the surface; preventing it touching down by smoothly increasing CL as drag decreases V², thus maintaining constant lift until CLmax is reached. At this point, the aircraft can no longer be 'held off' and it gently sinks the short distance onto the runway, touching down in a nose-high attitude. The acceleration or accelerated stall It is misleading to talk about stalling speed without further definition. The stall occurs at a particular aoa, not a particular speed. The speed — Vs — below which the stall will occur depends on the load factor. If the aircraft reaches the critical aoa under a load higher than 1g, the stalling speed will be higher than the normal 1g stall speed, at that mass. This latter stall is called an acceleration stall or accelerated stall and is usually more pronounced than a normal stall. The load factor normally increases in a turn— as we saw in section 1.10 where we calculated that, in a 45° banked turn, the load factor was 1.41 times normal. Thus, when turning, the stalling speed is higher than normal and the pilot must maintain a reasonable airspeed margin above that accelerated stall speed throughout the turn. See the table below. Be aware that the airspeed at which an acceleration stall in a turn occurs is only indirectly associated with the angle of bank; it is directly brought about by the increase in load factor. Indeed, it is possible to have the aircraft banked at 60° with a stall speed less than Vs1 if the wings are 'unloaded': slight forward pressure on the control column, and the aircraft allowed to sink, produces a load less than 1g — maybe 0.8g — with a stall speed less than Vs1, even though the aircraft is steeply banked. However, once the 'unloaded' condition ceases — if the stalling angle of attack has been passed (either by the rearward movement of the control column or a gust momentarily changes the relative airflow) — the probability of a stall returns immediately. The speed at which an accelerated stall occurs is proportional to the square root of the load factor — in the lift equation the airspeed is squared. If that load factor is expressed relative to the normal load, e.g. 2g, then the stall speed at that load factor — Vs 2g — equals the square root of the load factor × normal 1g stall speed; e.g. square root of 2 = 1.41 × Vs. The aircraft's momentum may also contribute to an accelerated stall, particularly when the aircraft is diving at speed and the pilot applies a harsh rearward control column movement. This will have the initial effect of rotating the aircraft about its lateral axis while inertia momentarily maintains the aircraft on its pre-existing flight path; thus the aoa may exceed the stalling aoa (even though the control column has not been pulled back to the normal stall position) with a consequent, and rather violent, high-speed stall. An acceleration stall can also be produced when: • the control column is jerked back whilst the aircraft is climbing or in level flight; see the flick roll • an aircraft in level cruising flight encounters a strong vertical gust • an abrupt change in flight path is made, which applies acceleration loads • an excessive bank angle, coupled with excessive control column back pressure, is applied during a level, climbing or descending turn. Note: The US Federal Aviation Regulations Section 23.203 airworthiness standards define accelerated stalls somewhat differently from the above, only referring to 'turning flight stalls' and 'accelerated turning stalls' for airworthiness demonstrations. "Turning flight and accelerated turning stalls must be demonstrated in tests as follows: (a) Establish and maintain a coordinated turn in a 30 degree bank. Reduce speed by steadily and progressively tightening the turn with the elevator until the airplane is stalled. The rate of speed reduction must be constant, and-- (1) For a turning flight stall, may not exceed one knot per second; and (2) For an accelerated turning stall, be 3 to 5 knots per second with steadily increasing normal acceleration*." * 'Normal acceleration' refers to the aerodynamic force parallel to the aircraft's normal axis. Load factor in a turn The table below shows the increase in stall speed at various bank angles in correctly executed level turns. The load factor or 'g' = 1/cosine of the bank angle and the Vs multiplier = the square root of the load factor. The table shows that once you reach bank angles of 30° or more, the aircraft stall speed increases rapidly; there is a 7% increase at 30°, 19% at 45° and 41% at 60°. Thus, level turns involving bank angles exceeding 20–30° should not be made at low levels, including take-off and landing operations. Even so, the airspeed should be increased to allow an appropriate safety margin — for gentle turns, a safe speed near the ground is 1.5 × Vs. The stall speed in a turn = Vsturn = Vs × Vs multiplier. A minimum turning speed at a safe height might be 1.2 × Vsturn. For example, if Vs is 50 knots and the bank angle is 45° then Vsturn is 50 × 1.19 = 60 knots and the minimum safe turning speed at height is 1.2 × 60 = 72 knots, or about 1.45 × Vs. Bank angle Cosine Load factor [g] Vs multiplier 10° 0.98 1.02 1.01 [+1%] 20° 0.94 1.06 1.03 [+3%] 30° 0.87 1.15 1.07 [+7%] 40° 0.77 1.30 1.14 [+14%] 45° 0.71 1.41 1.19 [+19%] 50° 0.64 1.56 1.25 [+25%] 54° 0.59 1.7 1.3 [+30%] 60° 0.50 2.00 1.41 [+41%] 70° 0.34 2.94 1.71 [+71%] 75° 0.25 4.00 2.00 [+100%] Note that the stall speed increases exponentially with bank angle; i.e. the 10° increase in the bank angles between 20° and 30° increases stall speed by another 4%, while the 10° increase in the bank angle between 50° and 60° increases stall speed by a further 16% (i.e. four times as much), while between 60° and 70° the stall speed is increased by a further 30%. Consequently aircraft certificated in the normal category are limited to a turning angle of bank of not more than 60°. Note that at an approach speed of 1.3 × Vs the aircraft will stall if turning with a 54° bank. The limits on climbing and descending turns are discussed in the 'Safety: control loss in turns' module. The torque stall For high-performance aircraft, with a very high power-to-weight ratio, the possibility of a torque stall exists. The most likely scenario is a sudden application of full power in a 'go-around' following an aborted landing, where the airspeed has been allowed to decay below the safety speed. The torque of the engine and inertia of the heavy propeller tends to twist the aircraft around the propeller shaft, and the consequent roll may increase the aoa of the downgoing wing past the critical aoa. If that happens, the wing loses lift, which accelerates the roll and the aircraft loses height very rapidly. However, torque stalls are probably not applicable to light aircraft, although the torque effect may influence the characteristics of a stall in a climbing turn. Effect of weight If the aircraft is below its MTOW, the operating wing loading will be less than the design W/S and the stall will occur at a lower speed than that marked on the ASI. For example, if we refer to the Jabiru, the wing area is 7.9 m², MTOW is 4200 N, Vso is 40 knots CAS and we can calculate that CL with flaps fully extended is 2.0. We saw above in the section 'The acceleration or accelerated stall' that W/S at the stall = CL × ½rV². We will rearrange that and say Vs² = (W/S) / (CLmax × ½r). Substituting the values, including 1.225 for density, we get: Vs² = (4200/7.9) /(2.0 × 0.5 × 1.225) = 532/1.225 = 434 m/s and Vs = 20.8 m/s = 40 knots CAS Now what will Vs be when the Jabiru with no passenger on board is at the low weight of 3400 N? Well, substituting that weight we get: Vs² = (3400/7.9) /(2.0 × 0.5 × 1.225) = 430/1.225 = 351 m/s and Vs = 18.7 m/s = 36 knots CAS. There are other, somewhat simpler, ways to calculate the reduction in Vs corresponding to a reduction in weight but what we see above is that a reduction in weight of 800 N, or about 19%, reduces Vs by 4 knots, or about 10%. This brings us to the mathematical rule of thumb that when two values are not that far apart in percentage terms, say up to 40%, their square roots are about half that distance apart in percentage; and because aerodynamic pressure is proportional to V², there are many occasions where the square root of a value is relevant. This allows a simple, but reasonably accurate, mental calculation: Rule of thumb #3 "The percentage reduction in Vs is half the percentage reduction in weight." i.e. If weight is reduced by 10% from MTOW then Vs will be reduced by 5%, and conversely, if weight is 10% over MTOW then Vs will be 5% higher — one of several reasons to avoid overloading an aircraft. (There is further discussion on weight control throughout these notes.) Thus in the section 'The acceleration or accelerated stall' above, where we referred to unloading the wings with the aircraft banked at 60°, the load reduction from 1g down to 0.8g is 20% so the unloaded stall speed would be about 90% of Vs1. You can also see the same relationship in the preceding table; for bank angles up to 45° the percentage increase in Vs is about half the percentage increase in W/S. It is appropriate to mention here that it is not only aircraft weight/wing loading that affects the stall speed. Some of the other critical performance values are also achieved at a particular aoa, and the associated airspeeds are also changed by a change in weight. The same rule of thumb applies to them. These critical performance values (the 'numbers') are: best rate of climb speed, best angle of climb speed, lowest power-off sink rate speed, best glide ratio speed and manoeuvring speed. Another aspect we will look at in the 'Aerofoils and wings' module is the effect of flaps, but we will just state here that flaps provide an increased CL at all angles of attack consequently allowing a reduction in V² and the stalling speed. In some aircraft extending flaps also increases wing area, thus W/S is reduced, a handy technique for high-performance military aircraft, manoeuvring at maximum allowed wing loading — they can tighten the turn even further without breaking the aircraft. The final aspect of the stall is the effect of atmospheric turbulence on aoa and this affects 'manoeuvring speed'. We will look at it in the 'Wind shear and turbulence" tutorial. 3.3.9 V-speeds Airspeed codes It is important to have a simple, easily understood and universally accepted identification method for the various airspeeds at which an aircraft may be operated, but currently it's a bit messy and there is no complete, unambiguous, and universally recognised, airspeed designation system published by any regulatory authority. Current nomenclatures are generally made up of two to six letters/numbers, with the first being V. Some of these V-speed codes — applicable to single-engine aircraft — with alternatives and definitions are shown below. These are relevant to sport and recreational aircraft including low momentum ultralight aircraft, and might appear in flight manuals, pilot's operating handbooks and even sales literature but those indicated with open bulleting º are probably only applicable to a few very light aircraft types. There are two classes of airspeed codes. One is the structural design speeds (specified in national airworthiness requirements) used in determining the airframe and control surfaces strength requirements for type certification. Such speeds include the term 'design' in their description. The other class are the designer recommended operating speeds. Please be aware that the various 'best' performance speeds mentioned below — rate of climb, angle of climb, cruising range, gliding range, etc. — merely indicate the midpoint in an airspeed range extending perhaps 1–2% either side of that point. Also, the performance speeds are very much affected by the horsepower of the particular engine fitted, plus the type of propeller and its pitch setting. If there is no pilot's operating handbook for the particular airframe/engine/propeller configuration, then the pilot must calculate the performance speeds by trial and measurement. Critical limiting speeds • Va — design manoeuvring speed. Design rules state that the minimum acceptable design manoeuvring speed is a fixed calculation relative to Vs1 for all aircraft within the same category. For a 'normal' category light aircraft (whose certificated load limit factor in the pitching plane is +3.8g), minimum Va = Ö3.8 Vs1, or 1.95 × Vs1. For a 'LSA' category aircraft (whose certificated vertical load limit factor is +4g), minimum Va = Ö4 Vs1, or 2 × Vs1. If the designer has opted for a design manoeuvring speed that is greater than the minimum acceptable speed then a Vo operating manoeuvring speed must also be specified. Va is also known as the 'optimum manoeuvre speed', or the 'corner speed', to military pilots as it's at the intersection of the structural limit load factor and the maximum aerodynamic force curve (the 'A' corner) in the aircraft's manoeuvring flight envelope, i.e. Va is the speed at which an aircraft can make the tightest possible turn (the minimum radius turn) and the fastest rate of turn by applying the aerodynamic limit (maximum aoa [CL max]) and the structural load limit simultaneously. (Of course, in the military context, in such a turn the aircraft would be comparatively low in kinetic energy.) In the sailplane context the symbol Vm is used for a manoeuvring speed which is the product of the square root of the design load limit factor and the minimum flight speed. Va is sometimes referred to as the 'speed for maximum control deflection' which has been the cause of much confusion. It is unwise to make full or abrupt applications of any one primary flight control if you are flying at a speed greater than Va, because at higher speeds it is easy to apply (see the stick force gradient) aerodynamic forces that could exceed the aircraft's structural limitations. But, even when flying below Va, it is unwise to make rapid control reversals or 'checks' such as alternating heavy applications of rudder or suddenly apply heavy asymmetric loads, e.g. heavy application of elevator and rudder or aileron; see the flick roll. (That misleading term, 'speed for maximum control deflection', was subject to much debate in 2001 following the horrific crash of American Airlines Flight 587, an Airbus A300 which, shortly after take-off while in a climbing turn at an airspeed 20 knots below standard Va, ran into wake vortices from a Boeing 747 four miles ahead. To counter sideslip it appears the pilot flying employed four nearly full-rudder reversal movements within a seven-second period. Those pilot-commanded side forces induced a tail fin load twice the design load limit and 1.3 times the ultimate load limit, at which point the complete fin and rudder broke away. In 2010 the U.S. Federal Aviation Administation issued a 'Maneuvering Speed Limitation Statement' namely: (i) Full application of pitch, roll, or yaw controls should be confined to speeds below the maneuvering speed; (ii) Rapid and large alternating control inputs, especially in combination with large changes in pitch, roll, or yaw, and full control inputs in more than one axis at the same time, should be avoided as they may result in structural failures at any speed, including below the maneuvering speed. ) Va is usually not marked on the ASI but there should be a placard indicating the MTOW manoeuvring speed on the instrument panel near the ASI or in the Pilot Operating Handbook or Aircraft Flight Manual; if not available, you can assume it's twice maximum weight Vs1 for non-aerobatic light aircraft and reduces as aircraft inertia (i.e. weight or, more properly, mass) reduces and thus Vs1 decreases. Note: if a recreational pilot is foolish enough to operate an aircraft at a weight exceeding MTOW then the overloaded aircraft's stall speed will be higher; consequently Va in that condition will be higher. So Va is not a fixed documented value, it decreases as the aircraft's weight decreases from MTOW, because the effects of the atmospheric forces become more pronounced as its inertia decreases. Sometimes the aircraft's documentation will specify the Va for weights below MTOW but it may be left up to the pilot to calculate the Va for the current aircraft weight. Using rule of thumb #3 above, the reduction in Va will be half the percentage reduction in aircraft weight; for example if, with only the pilot on board, weight is 16% below MTOW then Va is reduced by 8%. Flying at speeds below that estimated Va in turbulent conditions also enhances the possibility of stalls induced by vertical gusts — and also may reduce aileron and rudder effectiveness, so the pilot in command must be careful to select an optimum speed for the atmospheric conditions. Also note that the documented Va is calculated for the aircraft in a clean configuration — it does not apply to flight with the flaps extended, see Vfe maximum flaps extended speed. Misuse of controls in light aircraft can generate structural loads greater than those encountered in turbulence, so Va is also useful as a 'turbulent air operating speed' and most recreational aircraft operating handbooks recommend that airspeed be reduced to Va in turbulent conditions. When flying above this speed, gust-induced loads can exceed the design limit of many structures within the aircraft, including the pilot's seat and the engine mount, possibly even the crankshaft. Gust loads in the high temperature conditions of the Australian tropical continental air mass can be extremely high. Va is the recommended indicated cruising speed (CAS) when flying in moderate turbulence — intermittent, uncomfortable jolts. At this compromise speed, the aircraft will generally produce an accelerated stall and thus alleviate the aerodynamic force (including any manoeuvring forces) on the wings and other structures, if it encounters a vertical current that imparts an acceleration sufficient to exceed the load limit factor. Read 'The speed to fly in turbulence', in the 'Decreasing your exposure to risk' guide. • Vo — operating manoeuvring speed. If the aircraft designer has specified a design manoeuvring speed that is greater than the regulatory minimum (Ön × Vs1 where 'n' is the category limit load factor) then, when flying at Va and if a substantial nose-up pitching manoeuvre is applied, the aircraft may exceed the limit load factor before stalling. So, an operating manoeuvring speed Vo should also be established as an operating limitation speed, which is a selected speed that is not greater than Ön × Vs1 and is a speed where the aircraft will stall in a nose-up pitching manoeuvre before exceeding the structural load limits. Thus Va must be equal to or greater than Ön × Vs1 while Vo must be equal to or less than Ön × Vs1. The load limit factor is 3.8g for normal category sport and recreational aeroplanes and 4g for the LSA category. The square roots of 3.8 and 4 are very close — 1.95 and 2 respectively — so we can rephrase the preceding statement as 'Va must be equal to, or greater than, twice Vs1 while Vo must be equal to, or less than, twice Vs1. Remember that Vs1 stall speed reduces when aircraft weight is less than MTOW so the appropriate Va indicated air speed will decrease in proportion to the decrease in aircraft weight. Flying at speeds much below Va in turbulent conditions also enhances the possibility of stalls induced by vertical gusts — and also may reduce aileron and rudder effectiveness. • Vb — design speed for maximum gust intensity. The applicable vertical gust intensities range from 25 fps (7.5 m/s) to 50 fps (15 m/s). Also known as the maximum rough air speed. It is not required to be specified for normal, utility and LSA aeroplanes (in those categories Vb would generally only differ by a few knots from Va). Vb is specified in the European Joint Airworthiness Requirements JAR-22 for sailplanes and powered sailplanes in the utility and acrobatic categories but in this case Vb is the speed at which the sailplane is able to withstand a strong vertical gust of 50 fps (15 m/s or 30 knots) without exceeding the load limit factor, i.e. it is the speed at which an encounter with a gust of the specified value produces CL max. • Vfe — maximum flaps extended speed. It is indicated by the top end of the ASI white arc. Flight with flaps extended — or extending flaps — above this speed may result in distortion of the flaps or their supporting structure and extension mechanism. Various Vfe speeds may be specified according to the available flap settings. Generally speaking, the flight load limit factor is reduced by about 50% when flaps are fully extended. For example, the aircraft flight manual of a 'normal' category light aircraft will probably note that the load limit factor is reduced from 3.8g to 2g. So, extending some flap in turbulent conditions will decrease the stall speed but will reduce the load limit factor; thus it's very much a 'pilot-in-command' decision of 'when' and 'how much' to use, but certainly full flap is most unlikely in any condition. • Vno — maximum structural cruise speed or 'normal operating limit', indicated by the top end of the ASI green arc. Flight above Vno should only be conducted cautiously and in smooth air, while the pilot should not apply any abnormal control inputs when cruising at, or above, Vno. Vno must be equal to or greater than Vc (below, in section 'Cruise speeds'), but in most light aircraft Vno and Vc are assumed synonymous. When cruising at, or below, Vno, the aircraft should not be damaged by a 30 feet/second vertical gust — which is at the bottom end of the moderate to strong vertical gust scale of 25–50 feet/second vertical gusts; read 'The speed to fly in turbulence'. • Vne — never exceed speed, which is the IAS that should never be intentionally exceeded in a dive, or other manoeuvre in smooth air. FAR 23 requires that Vne be not more than 90% of the design diving speed Vd or the flight-demonstrated diving speed Vdf. Vd and Vdf are not included in pilot operating documentation — they are the realm of the test pilot. Vne is indicated by the red line at the top end of the ASI yellow arc. For light aircraft operating below 10 000 feet, it can usually be assumed that Vne is a fixed IAS. If aircraft have high altitude capability or particular airframe vibration characteristics, it is possible that the designer will specify Vne as a TAS, above a particular altitude and for various altitude bands of perhaps 3000 feet. If Vne varies with altitude, then FAR Part 23.1545 (c) requires a placard next to the ASI indicating the appropriate IAS limitations throughout the aircraft's operating altitude range. This particularly applies to sailplanes and powered sailplanes whose very high aspect ratio wings are candidates for flutter problems and such low drag aircraft can build up airspeed at altitude very quickly in quite shallow dives. For expanded information see 'Don't fly real fast' in the 'Decreasing your exposure to risk' guide. • Vs1 (sometimes incorrectly shown as Vsi) — stalling speed, or the minimum steady flight speed, in a specified flight configuration. For a simple aircraft, Vs1 is normally measured in level flight with flaps up, at MTOW and 1g wing loading, with engine idling following a gradual deceleration (one knot per second) — accompanied by increasing rearward movement of the control column — to that minimum flight speed. It is indicated by the bottom end of the ASI green arc, but it should be documented as both IAS and CAS; if CAS is not mentioned the quoted stall speed is probably inaccurate. Vs1 decreases as the aircraft weight decreases from MTOW, which also means that if the pilot can reduce the wing loading below 1g — by an 'unloading' manoeuvre — Vs1 is decreased. Stalling speed under a 2g wing loading, for instance, might be referred to as Vs2g. • Vso — stalling speed, or the minimum steady flight speed, in the landing configuration of flaps down and engine at low or idle power as it would be just prior to touchdown. This is measured using the same method as Vs1 but with the cg at the most extreme position allowed — usually the most forward position where backward movement of the control column may be limited. It is indicated by the bottom end of the ASI white arc.In the documentation both IAS and CAS should be shown. Like Vs1, Vso decreases as the aircraft weight decreases from MTOW. The designation Vs is used as a general reference to the design stall speed. • Vmin — minimum airspeed. Vs is generally specified in powered, rigid-wing recreational aircraft as the minimum speed but for other aircraft categories a 'Vmin' may be specified instead. For example for gyroplanes Vmin is the minimum controllable level flight airspeed below which there is insufficient power available to maintain altitude. For paragliders it is the minimum speed, within the wing's available trimmer range, below which the parawing loses its lift. Cruise speeds A cruising aircraft is normally flying at a moderate, fuel efficient speed and maintaining the appropriate cruising altitude. The Australian Civil Aviation Regulations hold this definition: "cruise phase of flight" means the part of an aircraft's flight: (a) that starts when the aircraft reaches its first planned cruise level, ... and (b) that ends when the aircraft reaches the point at which the aircraft first starts its descent for the purpose of landing; and includes flight level changes made during that part of the flight. • Vbr — best range, or Vmd — minimum drag, is the speed that provides maximum L/D by producing minimum drag and thus the best power-to-speed ratio. This speed might utilise about 55% power and is usually flown at the lowest altitude where the throttle is fully open to obtain that speed. Vbr/Vmd decreases as the aircraft weight decreases from MTOW. It's rather boring to fly at that speed, wind conditions have to be taken into account, and the fuel saving may not be that significant compared to flying at a speed 10% faster. Also, the engine manufacturer's operating recommendations should be followed, but mixture is usually leaned, and minimum rpm set if a constant speed propeller is fitted. Vbr/Vmd has the same basic airspeed range as Vy and Vbg [below]. There is a difference in concept between Vbr and Vmd. Pilots of low-powered aircraft are generally not interested in the best power to airspeed ratio in cruise; ground speed achieved per litre is far more significant, so in some conditions Vbr equals Vmd but in headwind conditions Vbr is increased. Look at the diagram from section 1.7 at left and note the pink line that has been drawn from the junction of the vertical and horizontal axes tangential to the power required curve. That line just touches the curve at a position corresponding to the minimum drag airspeed Vmd. Now imagine a 30 knot headwind and start the tangential line from a point along the horizontal axis that is equivalent to 30 knots; that (the blue line) will be tangent to the power required curve at a position corresponding to a higher speed — Vbr for a 30 knot headwind. The rule of thumb is to add half the head wind to the basic Vbr, which, in this case, indicates a Vbr that is about 15 knots greater than Vmd. This is the same principle used by sailplane pilots to establish their best penetration speed — see the speed polar curves for optimum glide speed in the 'Coping with emergencies guide'. • Vbe — best endurance, or Vmp — minimum power, is the CAS that gives the greatest airborne time per litre (i.e. least fuel flow per hour and, of course, power is proportional to fuel flow), possibly around 80% of Vbr/Vmd, and decreases as the aircraft weight decreases from MTOW. Flight at lowest safe altitude provides best engine performance. Might utilise about 45% power at MTOW. It is the speed for minimum power required for level flight, as shown in the power required curve above. Vbe/Vmp is the speed that might be used when flying a search pattern to allow a proper area survey, or when waiting for ground fog to disperse, but it is possibly uncomfortable to fly for long periods at such a low speed. Also the very low power setting may be inconsistent with good engine handling practice. Carburettor icing may be aggravated. The Vmp designation and speed is also used as a power-off glide speed, providing the best endurance — least rate of sink — in the glide; see 'Power-off descent speeds' below. Vbe/Vmp is in the same speed range as Vx — the best angle of climb airspeed. • Vc — the design cruising speed or the optimum cruise speed — the latter being the speed that gives the most velocity (i.e. greatest distance/time) from a litre of fuel, usually utilises 75% power at MTOW and is about 20–30% greater than the maximum L/D speed — Vbr. The speed and power required both decrease as the aircraft weight decreases from MTOW. Refer to rule of thumb #3 in section 2.8 'Stalling airspeeds'. For normal category aircraft, FAR Part 23 specifies a minimum design cruising speed (in knots) = 33 ÖW/S. For this calculation the wing loading W/S is expressed in pounds per square foot. Many minimum ultralights are unable to comply with the FAR Part 23 design requirement for a minimum design cruising speed. For most light aircraft, Vc is synonymous with Vno. FAR 23 Appendix A provides simplified design load criteria and allows designers of many conventional single-engine monoplanes weighing less than 2700 kg to take advantage of the simplification. That same appendix is generally duplicated in the design regulations of most other countries. One advantage is that it is not necessary to specify Vno; instead, Vc is designated in the flight manual as the maximum structural cruise speed (i.e. Vno = Vc) and that Vc may be set at 90% of Vh. • Vh — the maximum level flight indicated speed (CAS) attainable at sea-level, utilising maximum continuous engine power. For most engines maximum continuous engine power at sea-level will be less than full throttle power. Take-off and landing speeds • Vle — for retractable undercarriage aircraft — the maximum indicated speed at which the landing gear can remain extended without risking gear door damage. • Vlo — the maximum indicated speed at which the landing gear system can be operated. Vle and Vlo are unlikely to be applicable to most ultralights. • Vlof — the lift-off indicated speed for normal take-off. Vlof is about 10% above Vmu. • Vmu — minimum unstick speed. This is an indicated speed used in take-off conditions where it is advisable to lift off at the lowest possible airspeed to get the tyres off the surface (e.g. soft field or wet grass ) and safely fly in ground effect until a Vtoss is attained to allow climb-out. Acceleration after lift-off at Vmu is slow, due to the drag at the high aoa, and should not be used as an obstacle clearance technique. • Vref — the threshold speed or the reference indicated approach speed. Usually about 1.3 to 1.5 times Vso plus 50% of the wind gust speed in excess of the mean wind speed; e.g. Vso = 30 knots, wind speed 10 knots gusting to 20 knots, Vref = 1.3 x 30 + 5 knots = 44 knots. Faster, heavier aircraft would tend towards the 1.3 times Vso end; lighter, slower aircraft would tend towards the 1.5 times Vso end. Normal landing procedure is to set up the approach so that an imaginary 15 metre (50 ft) high screen placed before the runway threshold is crossed at Vref and the airspeed is reduced to maybe 1.2 to 1.3 × Vso — plus the gust allowance — when rounding out prior to touchdown. The ground distance from the screen to the touch-down point can be roughly estimated, using the 1-in-60 rule, from the approach slope. For example, with a 6° slope — which is around the norm for most light 3-axis aircraft — the distance will be 60/6 × 15 = 150 m. To this must be added any float period plus the ground roll distance with normal braking, to give the total landing distance over the standard 15 m screen — in nil wind conditions. • Vtoss — minimum take-off safety speed. This is an indicated speed chosen to ensure that adequate control will still exist during initial climb after lift-off if power is lost or turbulence encountered. After lift-off, the aircraft should be held down and not allowed to climb away until Vtoss is attained. CAO 101.28, an airworthiness certification requirement for commercially supplied amateur-built kit ultralights, states in part: "The take-off distance shall be established and shall be the distance required to reach a screen height of 50 feet from a standing start, with ... short, dry grass surface ... the aeroplane reaching the screen height at a take-off safety speed not less than 1.2 Vs1 ... take-off charts ... shall schedule distances established in accordance with the provisions of this paragraph, factored by 1.15." Sea-level ISA and nil wind conditions are implied. CAO 95.55 has much the same wording but specifies 1.3 Vs1 as the take-off safety speed. (Similarly, CAO 101.28 states that the landing distance will be that to come to a full stop from a screen height of 50 feet, with the screen being crossed at 1.3 Vso and the same conditions as specified for the take-off distance.) In normal take-off conditions Vtoss should be somewhere between 1.3 and 1.5 times Vs1, with 'draggy' aircraft tending to the higher value. If power is lost in the initial climb, a draggy aircraft will lose airspeed very rapidly and take some time to regain it even though the pilot reacts quickly and pushes the control column forward. See 'Engine failure after take-off'. There is a similar code used for multi-engine aircraft — Vtos — which refers to the minimum speed for climb-out with one engine inoperative. Climb speeds • Vx — indicated speed provides best angle of climb for obstacle clearance; i.e. to attain height over the shortest ground distance using maximum thrust available. This is probably better described as the precautionary climb speed — the initial climb speed used when there are obstructions off the end of a marginal airstrip or when climbing out of an obstructed valley. Vx decreases as the aircraft weight decreases from MTOW (refer rule of thumb #3 above), but the angle of attack is maintained at around 8–10º with very high induced drag. It is the climb airspeed where the ratio of vertical speed to horizontal (ground) speed is the highest. Vx may be less than or equal to Vtoss. The aircraft's power-to-weight ratio (i.e. excess power) and L/D ratio affect the angle of climb at the designated airspeed. However, be aware that the angle of climb will also depend on the low-level wind conditions at the airfield. In a headwind, the climb angle is increased and reduced in a tailwind. Also note that aoa during climb may be only 5 or 6° below the critical aoa, thus care must be taken not to induce a 'departure stall', particularly in turbulent conditions. And remember that Vs1 increases in a turn, so that the small safety gap between Vx and Vs1 will be eroded if a climbing turn is attempted; see 'Safety: loss of control in low level turns'. Climbing at Vx should always be regarded as a short-term precautionary procedure, and once clear of obstacles, airspeed should be increased to Vy — or any appropriate 'enroute climb speed'. The latter reduces the rate of climb but has the benefit of reducing total sector time, increasing forward visibility and increasing engine cooling — which may be beneficial to engine operation but, more importantly, provides a little more airspeed in hand should the engine falter or fail. The airspeed for Vx increases with (density) altitude and is much the same airspeed as Vbe, although engine cooling needs might require a higher airspeed. • Vy — indicated speed for best rate of climb. This speed is used to attain height in the shortest time using maximum power, or possibly maximum continuous climb power. Vy decreases as the aircraft weight decreases from MTOW (refer to rule of thumb #3 above), but the angle of attack is maintained at around 6–8º. After reaching a safe height airspeed may be increased to an appropriate enroute climb speed. The CAS for Vy decreases with (density) altitude — i.e. as TAS increases — and also is usually fairly close to the maximum L/D speed Vbr, taking engine cooling flows into account. Vx and Vy converge as (density) altitude increases. Power-off descent speeds • Vbg — best power-off glide This is the airspeed that provides minimum drag thus maximum L/D, or glide ratio, and thus the greatest still air glide range from the potential energy of height. It is much the same basic airspeed as Vbr/Vmd and Vy, though it may be a bit lower and decreases as the aircraft weight decreases from MTOW. However, like Vbr, wind direction and speed have to be taken into account before you can choose the Vbg speed when in a forced glide; for more information on the power-off glide speeds read the 'Know the best glide and minimum descent airspeeds' and 'Know the practical glide ratio and terrain footprint' sections in the 'Coping with emergencies guide'. In lower wind conditions, Vbg is increased in a headwind by around one quarter of the windspeed, but is decreased in a tailwind by a similar amount. In higher wind conditions, say above 25 knots, the speed changes required would be around one half of the windspeed. • Vmp — minimum power. This is the speed that results in the lowest rate of sink in a power-off glide, and provides the longest duration of flight from the potential energy of height. The lowest rate of sink occurs at the minimum value of drag × velocity. It is probably around 80–85% of Vbg, and may be a similar speed to Vbe and Vx. Vbg for an average sailplane with a wing loading of 32 kg/m² could be 50 knots, providing a glide ratio of 38:1, while Vmp would be 41 knots providing a sink rate of 0.6 m/s or 120 feet/minute. If you want further explanation of sink rates, etc. (with excellent diagrams) read this article on glider performance airspeeds. (Note: the term Vmd meaning minimum descent, rather than minimum drag, is in common usage to designate the speed for lowest rate of sink in a power-off glide.) Both Vbg distance and Vmp time are adversely affected by the extra drag of a windmilling propeller, which creates more drag than a stopped (but unfeathered) propeller following engine shut-down or failure. A windmilling propeller has a negative aoa and the 'thrust' direction is reversed, in effect adding to drag. Something similar might happen with some engine/propeller configurations, when simulating a glide at the specified Vbg/Vmp speed with the throttle closed but the engine still firing; the propeller drag might increase the rate of sink beyond that expected, and perhaps lead to the erroneous conclusion that the best glide speeds in the handbook are understated. 3.3.10 The design manoeuvring flight envelope The structural design manoeuvring flight envelope of a recreational aircraft describes its structural strength limitations and its maximum aerodynamic capabilities. The various envelopes have been described as "the parameters within which an aircraft can be safely operated, with average pilot ability, at varying density altitudes, airframe states, power outputs, wing loadings and atmospheric turbulence". 'Airframe states' refer to flap and spoiler extensions, undercarriage position, the gross weight and the fore-and-aft position of the centre of gravity. The flight envelope is only relevant for an aircraft within the required weight and balance conditions. The boundary of the envelope shown below is formed by combinations of the applicable limiting positive and negative load factors and dynamic pressures (indicated airspeeds). It is a two-dimensional model that has indicated airspeed (i.e. CAS) along the horizontal axis and flight load factors expressed in 'g' accelerations along the vertical. The loads are those parallel to the aircraft's normal axis, i.e. perpendicular to the longitudinal axis — the aerodynamic loads in the aircraft's pitching plane. The symbol 'n' is generally used to identify such loads. The parameters for a light aircraft usually are the limiting critical airspeeds — Vs, Va and Vne and the certificated limit load factors. There are other flight limitations which are not displayed in this flight envelope diagram, e.g. an angle of bank limitation of 60° for some aircraft categories. For weight-shift aircraft in particular, there are also pitch limitations; — e.g. 45° nose up or nose down from the horizontal. The manufacturers of high performance non-recreational aircraft would also provide other charts, altitude performance or turning performance for example. The V-n [or V-g] diagram below is a simplified representation of a few aspects of the manoeuvring flight envelope for an LSA category aircraft at MTOW and low altitudes. An indicated (CAS) airspeed scale would normally be displayed along the horizontal axis and load factors (in units of 'g') along the vertical axis, between the certificated load limits for light sport aircraft of +4g to –2g. This diagram does not reflect any flaps-extended conditions. The positive and negative 'aircraft normal force coefficient curves'* are for the aircraft as a whole but can be assumed to approximate the accelerated stall speeds at loads of Vs1g × Öload factor. (*Normal force: pilots consider the aerodynamic force acting on the aircraft in terms of lift and drag, lift being perpendicular to the relative free-stream airflow and drag being parallel with it. Aerodynamicists may also consider the total aerodynamic force in terms of a 'normal force' which is parallel to the aircraft's normal axis and an 'axial force' which is parallel to the longitudinal axis plus a 'side force' parallel to the lateral axis. The symbol for the aircraft normal force coefficient is Cna, however, the stated aircraft limit load factor (e.g. 3g or 4g) also applies to axial and side forces, not just the normal force.) The stall speed at a 4g load limit factor would be Vs1g × Ö4 = Vs1g × 2, thus if the Vs1 stall speed was 45 knots then the stall speed where the positive curve intersects the 4g load limit factor line is 90 knots, i.e. that speed is the lowest possible speed at which the pilot can pull maximum g, i.e. 4g. That corner of the envelope is usually the position of Va — the design manoeuvring speed. You can see from the curve that at the Va airspeed the aircraft will stall when the wing loading exceeds 4g. Sustained flight is not possible in the white region to the left of the accelerated stall curves because the wings will be stalled. (Note: the light blue area between +1g and –1g is the realm of reduced gravity, or microgravity. NASA and other organisations use C135 and DC-9 aircraft flying a parabolic trajectory to produce reduced or near-zero gravity conditions, for the aircraft occupants, for periods of 20–30 seconds. A light aircraft can be flown in that area for a brief period by 'unloading' the wings — 'bunting'.) Except for transient turbulence loads the negative flight envelope below the 0g line mainly relates to aerobatic aircraft. (Note: in section 2.8 we determined that a 60° banked level turn doubled the normal wing load. With Vs1 at 45 knots and Va 90 knots then visualise a horizontal line from the 2g point; the interception with the curve will equate with about 60 knots. So that would be the lowest possible speed for a 60° banked level turn.) The white areas above and below the red lines represent structural loads beyond the positive and negative limit loads. Flight loads caused by control misuse and/or atmospheric turbulence that exceed +4g or –2g may cause temporary pilot incapacitation (greyout/blackout/redout) and airframe distortion. Flight loads 50% greater than +4g or –2g (i.e. +6g or –3g) will very likely cause airframe breakup. In the aircraft design process the design maximum dive speed Vd is a calculated speed, but in the flight test stage the aircraft may be tested up to a speed where it still demonstrates no flutter, or other, problems. This is the flight-demonstrated dive speed Vdf which is lower than the design Vd but, possibly, it could be equal to it. The dynamic pressure that a pilot must not exceed is represented by the Vne airspeed limit, and that is required to be no more than 90% of Vd or Vdf. At Vne the aircraft is flying at a very small angle of attack, deriving most of the aerodynamic force from the dynamic pressure. If the pilot — or turbulence — suddenly increases the aoa the consequent increase in the lift coefficient CL (amplified by the aircraft's inertia momentarily maintaining the original flight path) could place an extreme load on the airframe, enough to break it. See 'Wind shear and turbulence'. Some manoeuvring flight envelopes might have the top right corner cut off, from the Vno line intersection with the limit load factor line, to some less-than-maximum load factor along the Vne line. For more information concerning the risks of flight at excessive speed read 'Don't fly real fast!' in the 'Decreasing your exposure to risk' tutorial. Vne is the maximum airspeed, but full and rapid control applications are restricted to speeds at or below Va. Vno is the maximum structural cruise speed or 'normal operating limit' for flight in light to medium turbulence. Above the Vno/Vc speed flight should only be conducted cautiously and in smooth air. So, the aircraft can be flown in the light green area without limits on smooth control use and it can be operated within the olive-green area in light to moderate turbulence, but it should not be operated in the yellow area except in a reasonably smooth atmosphere. If it is inadvertently operated in the area outside the certificated load limits, or at velocities greater than Vne, structural distortion then failure may result. The more the wings are loaded while the aircraft is operating in the region above Vne, the greater the possibility of structural failure. The potential exists to exceed both Vd and maximum load in the pullout from a spiral dive. Vertical gusts impose loads on the wing structure by inducing rapid, but momentary, changes in aoa with consequent changes in the aerodynamic forces. The faster the aircraft is moving, the greater the gust-induced load. FAR Part 23 has requirements for designers to consider unexpected gust loads. The resulting gust envelope is often represented as the flight manoeuvring envelope with overlaid gust lines. Vb is developed by the aircraft designer as a recommended turbulence penetration speed in severe turbulence, with varying vertical gust components — up to 50 feet/second considered for a light aircraft at cruise speed. However, Vb is not specified for most light aircraft because, for such aircraft, there is probably not much difference between Va and Vb. The flight envelope is considerably reduced if asymmetric manoeuvring loads are applied to the airframe. Such loads might be applied by an aircraft yawing (side force) or rolling (lateral force) while recovering (normal force) from a high-speed descent. The same applies to the use of flaps. There are other attributes that define the envelope – resistance to spin and spin recovery, for example. Note that the term 'average pilot ability' doesn't imply that those who consider themselves 'above average' can push the envelope without losing control or stressing the airframe. There is more information on the flight envelope in the safety brief document 'Don't fly real fast'. Things that are handy to know • Absolute temperature is expressed in kelvins [K], one K equals 1 °C. The base temperature is zero kelvin — equivalent to minus 273 °C — so 0 °C is equivalent to 273 K. • In a free stream airflow, a favourable pressure gradient is one where static pressure decreases with distance downstream. An adverse pressure gradient is one where static pressure increases with distance downstream. • ASI position error. The static vent is an opening, best placed at a position on the aircraft's fuselage, where atmospheric static pressure is not influenced by the shape of the fuselage or other aerodynamic disturbances. (Some aircraft may be fitted with a static vent on each side of the fuselage to counteract static pressure disturbances caused when the aircraft is slipping/skidding, and/or a switchable alternative static source within the cockpit.) The opening is a tube connected to the cockpit and supplies the ambient atmospheric pressure, or static pressure, to the three pressure sensing instruments — ASI, VSI and altimeter. The static vent is usually subject to some pressure disturbances at particular aircraft attitudes, as is the pitot tube, but probably to a lesser degree. In addition if the airflow is not squarely aligned with the entry of the pitot head there will be a reduction in the indicated airspeed which increases as aoa increases. These disturbances result in position error: for a wing-mounted vent, the ASI may underread by 10 knots at stalling aoa. In a sideslip, a single fuselage-mounted static vent may be subject to dynamic pressure and ASI and VSI readings will consequently be completely misleading. Also, the instrument movements will have inbuilt errors, usually caused by excessive friction. Obstructions in the tubes — such as water or wasp's nests — will cause misreadings or no reading. Position error corrections plus the instrument error corrections for the system should be noted in the Pilot's Operating Handbook and placarded on the instrument panel. The IAS corrected for instrument and position errors is called the calibrated airspeed [CAS]. Either CAS or IAS may be the reference speed in the Pilot's Operating Handbook for aircraft operations, but if the position error corrections are not shown then the ASI system has not been assessed for accuracy. In some poor ASI installations, IAS may be 20% less than CAS at low speeds, but they are usually much the same at normal cruising speeds. Regulations for type-certificated aircraft require that the complete airspeed indicating system of pitot head, static vent, connecting tubes and instrument be tuned so that the IAS reading is within 3% of the true reading over the normal airspeed range from Vs to Vc. However, you should suspect that any non-certificated ultralight ASI system will be inaccurate at all speeds, and particularly so at high aoa. When comparing published stall speeds between different aircraft types, it is wise to determine CAS, as published IAS stall speeds may be downright misleading. • Compressibility effects. The compressibility of air within the pitot tube has little effect on the accuracy of the ASI reading for aircraft operating below 10 000 feet and 200 knots; at an airspeed of 200 knots, compressibility will cause CAS to overread by only 0.5 knots or so. However, for aircraft operating at high speed or high altitude, compressibility will cause the ASI to overread significantly, so there is a need to correct CAS using a compressibility correction chart. The correction value is deducted from CAS to give the compressibility corrected CAS — otherwise known as equivalent airspeed [EAS]. For most medium-speed aircraft, it is probable that the compressibility correction value has been built into the IAS–CAS airspeed correction table. There is no practical application for recreational pilots, but aerodynamicists use the EAS term — rather than IAS or CAS — assuming an ASI, that has no errors caused by mechanical, position, aoa or compressibility effects, would display the ISA standard condition sea-level true airspeed, which is equivalent to the dynamic pressure in the instrument at any altitude. For more information see 'Notes: compressibility of airflow and Mach number'. • Checking validity of claimed stall speeds. There is a simple method to check the validity of published stall speeds. Practically all very light aircraft (except those with single surface wings like the Wheeler Scout or weight-shift aircraft) use simple, long proven, standard camber aerofoils to form the wings. The lift coefficient attainable at maximum aoa with such wings without flaps is about 1.2 or 1.3 for faster-cruising aircraft, and 1.5 or 1.6 for the slower, higher-lift sections. If equipped with flaps over, say, half the trailing edge, then CLmax might be increased by 0.5 when the flaps are extended to at least 35°. When other high-lift devices (for example, full length leading edge slats/slots) are added to the wing, then CLmax might increase 0.6. Thus, a specialised short take-off and landing aircraft fitted with a high-lift aerofoil, full-length leading edge slats and large extended flaps would have a CLmax of (at least) 1.6 + 0.5 + 0.6 = 2.7. The lift equation at normal stall speed is: Lift = CLmax × ½rV² × S = weight or re-arranged: CLmax = weight / (½rV² × S) We can use that equation to check the validity of stall speed claims if we know the maximum take-off weight [MTOW] and the wing area . Let's say a supplier claims that an aircraft, lacking any high-lift devices, has a stall speed of 30 knots. The MTOW is 450 kg and the wing area is 12 m². In the equation, the weight must be expressed in newtons — so multiply kg × 10 = 4500 N; and the stall speed must be expressed in metres per second — so just halve the velocity in knots = 15 m/s: the air density used must be the ISA msl density = 1.225 kg/m³. Thus CLmax = 4500 / (0.5 × 1.225 × 15 × 15 × 12) = 2.7 A lift coefficient of 2.7 is very much higher than that achievable without high-lift devices, so you would conclude that the claimed stall speed is nonsense; a figure of 38 knots is probably closer to the mark. • Estimating stall speeds. Conversely you can do a rough approximation of stall speeds using the following simplified formulae if you know the wing loading in kilograms per square metre or in pounds per square foot, and can estimate CLmax with flaps stowed or fully extended. Stall speed [knots] = 7.8 × square root (wing loading in kg/m² divided by CLmax) (or) Stall speed [knots] = 17.2 × square root (wing loading in lb/ft² divided by CLmax) Using our previous example of a lightly-loaded Jabiru with a mass of 340 kg (748 lb), wing area of 7.9 m² (85 ft²) thus wing loading = 43 kg/m² (8.8 lb/ft²) and estimating CLmax with flaps fully extended as 2.0 then: estimated stall speed = 7.8 × square root (43/2) = 7.8 × 4.64 = 36 knots or estimated stall speed = 17.2 × square root (8.8/2) = 17.2 × 2.1 = 36 knots Stuff you don't need to know • The tropopause marks the boundary between the two lower layers of the atmosphere — the troposphere, and above it, the stratosphere. The height of the tropopause varies daily, seasonally, and latitudinally — it is about 28 000 feet at the poles and perhaps 55 000 feet at the equator. The significant difference between the troposphere and the stratosphere is that air temperature decreases steadily with height in the troposphere, but initially remains constant then increases steadily with height in the stratosphere, until the stratopause at about 50 km. The stratosphere contains very little water vapour and is much more stable than the troposphere. The ozone layer is within the stratosphere. • Boyle's law: at a constant temperature, the volume (V) of a given mass of gas is inversely proportional to the pressure (P) upon the gas; i.e. PV = constant. • The pressure law: at a constant volume, the pressure is directly proportional to temperature (T) in kelvins. • Charles' law: at a constant pressure, gases expand by about 1/273 of their volume, at 273 K, for each one kelvin rise in temperature; i.e. the volume of a given mass of gas at constant pressure is directly proportional to the absolute temperature. • For one mole of gas, the preceding laws are combined in the gas equation PV = RT, where R = the gas constant = 2.87 when P is expressed in hectopascals. Ordinary gases do not behave exactly in accordance with the gas laws. • The change in altitude for each one hPa change in pressure can be roughly calculated from the absolute temperature and the pressure at the level using the equation:=96T/P feet. • The term 'burble' also refers to the atmospheric wake of an object. Skydivers refer to their wake as 'the burble' while the disturbed airflow and exhaust gases behind a the island structure of a fast-moving aircraft carrier was (and probably still is) known to pilots as 'the burble'. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.2.1 Cruise performance Cruise speed options When an aircraft is cruising, flying from point A to point B, the pilot has several options for setting cruise speed: One choice might be to get there as soon as possible, in which case the pilot would operate the engine at the maximum continuous power allowed by the engine designer. The recommended maximum continuous power is usually around 75% of the rated power of the engine and provides performance cruise. Another choice might be to get there using as little fuel as possible but in a reasonable time, in which case the pilot might choose a 55% power setting to provide an economy cruise airspeed. Or the pilot might choose any power setting, in the usual engine design range, between 55% and 75%; refer to cruise speeds in the 'Airspeed and properties of air' module. The power required curve In level flight at constant speed thrust power is required to balance induced and parasite drag. Power is the rate of doing work, so power (in watts) is force (newtons) × distance (metres) / time (seconds). Distance/time is speed so power required is drag force (N) × aircraft speed (m/s). Thus, if we use the expression for total drag from section 1.6 and multiply it by V we get: (Equation #1.3) Power required for level flight [watts] = CD × ½rV³ × S (note V cubed). The total drag curve can be converted into a 'power required' diagram — usually called the power curve — if you know the total drag at each airspeed between the minimum controllable speed and the maximum level flight speed. It is a different curve from that for total drag, because the power required is proportional to speed cubed rather than speed squared. This means that (ignoring the related CD change) if speed is doubled, drag is increased four-fold but power must be increased eight times — which indicates why increasing power output from, say, 75% power to full rated power, while holding level flight, doesn't provide a corresponding increase in airspeed. The diagram above is a typical level-flight power curve for a light aircraft. The part of the curve to the left of the minimum power airspeed is known as the back of the power curve — where the slower you want to fly, the more power is needed, because of induced drag at a high angle of attack. The lowest possible speed for controlled flight is the stall speed, which is discussed in the 'Airspeed and properties of air' module. Two aerodynamic cruise speeds are indicated — the speed associated with minimum drag (the point on the curve where the drag force factor has the lowest value) and the speed associated with minimum power (the point on the curve where drag force × speed has the lowest value). To maintain level flight at speeds less than or greater than the minimum power airspeed, power must be increased. Power available The engine provides power to the propeller. The propellers used in most light aircraft have a maximum efficiency factor, in the conversion of engine power to thrust power, of no more than 80%. (Thrust power = thrust × forward speed.) The pitch of the blades, the speed of rotation of the propeller and the forward speed of the aircraft all establish the angle of attack of the blades and the thrust delivered. The in-flight pitch of ultralight and light aircraft propeller blades is usually fixed (though many such types are adjustable on the ground) so that the maximum efficiency will occur at one combination of rpm and forward speed — this is usually in the mid-range between best rate of climb and the performance cruise airspeeds. Propeller blades are sometimes pitched to give the best efficiency near the best rate of climb speed (climb prop), or pitched for best efficiency at the performance cruise airspeed (cruise prop). The efficiency of all types of propellers falls off either side of the optimum; one with a too high pitch angle may have a very poor take-off performance, while one with a too low pitch may allow the engine to overspeed at any time. With the advent of higher-powered four-stroke light engines, such as the Jabiru 3300, there has been a corresponding increase in the availability of more advanced light-weight propeller systems, providing maximum effective power utilisation during all stages of flight. For more information refer to the 'Engine and propeller performance' module. Speed, power and altitude At sea-level, an aero-engine will deliver its rated power — provided it is in near-perfect ex-factory condition, properly warmed up and using fuel in appropriate condition. However, because air density decreases with increasing altitude, and an engine's performance depends on the weight of the charge delivered to the cylinders, then the full throttle power of a non-supercharged four-stroke engine will decrease with height. So, at about 6000–7000 feet above mean sea-level the maximum power available at full throttle may drop below 75% of rated power. At 12 000 feet full throttle power may be less than 55% of rated power. Thus, as altitude increases, the range of cruise power airspeeds decreases. For best engine performance, select a cruise altitude where the throttle is fully open and the engine is delivering 65% to 75% power. A couple of points to note from the speed-power diagram above: As air density, and consequently drag, decreases with height, then airspeed, from a particular power level, will increase with height; e.g. the airspeed attained with 65% power at sea-level is 90 knots increasing to 100 knots at 10 000 feet. At sea-level, an increase in power from 75% to 100% only results in an increase in airspeed from 100 to 110 knots. This is the norm with most light aircraft — that last 33% power increase to rated power only provides a 10% increase in airspeed. Power required vs power available In the 'power available' diagram at left, power available curves have been added to the earlier 'power required' diagram. The dashed red curve indicates the rated power — that is, the full throttle engine power delivered to the propeller over the range of level flight speeds at sea-level. The upper green curve — maximum thrust power, is that engine power converted by the propeller after allowing for 80% maximum propeller efficiency. The lower green curve is the propeller thrust power available with the engine throttled back to 75% power at sea-level, or if flying at an altitude such that full throttle opening will only deliver 75% of rated power. The intersection of those power available curves with the power required curve indicates the maximum cruise speed in each condition. The region between the maximum thrust power curve and the power required (to maintain level flight) curve indicates the excess power available at various cruise speeds — this excess power is available for various manoeuvres if the throttle is fully opened. The simplest use would be a straight unaccelerated climb, in which case the maximum rate of climb would be achieved at the airspeed where the two curves are furthest apart. It can be seen that the best rate of climb speed is around the same airspeed as the minimum drag airspeed shown in the earlier powered required diagram. The rate of climb will decrease at any speed either side of the best rate of climb speed because the power available for climb decreases. The rate of climb (metres/second) = excess power available (watts)/aircraft weight (N). For example, lets assume the preceding diagram is representative of an aircraft fitted with a 100 hp engine, and at the best rate of climb speed the engine/propeller has 25 hp (18 600 watts) of excess thrust power available. The aircraft weight is 4000 N so the rate of climb = 18 600/4000 = 4.65 m/s. To convert metres/second to feet/minute, multiply by 200 = 930 feet/minute as the maximum rate of climb. One thing to bear in mind is that we have assumed the aircraft's aerodynamic shape — its configuration is constant. However if the aircraft is fitted with flaps, high lift devices or spoilers the pilot is able to change its configuration and consequently its performance. Thus performance is dependent on power, plus attitude (pitch, bank, sideslip and aoa) plus configuration. 3.2.2 Forces in a climb When cruising, the difference between the current power requirement and power available — the excess power — can be used to accelerate the aircraft or climb, to accelerate and climb, or perform any manoeuvre that requires additional power. For instance if the aircraft has potential power available and the pilot opens the throttle, the thrust will exceed drag and the pilot can utilise that extra thrust to accelerate to a higher speed while maintaining level flight. Alternatively the pilot can opt to maintain the existing speed but use the extra thrust to climb to a higher altitude. The rate of climb (altitude gained per minute) depends on the amount of available power utilised for climbing, which depends in part on the airspeed chosen for the climb. There are other choices than the best rate of climb speed available for the climb — for example, the best angle of climb speed (which is around the same as the speed for minimum power) or a combination enroute cruise/climb speed. The climb speed chosen depends on terrain, weather, cloud cover and other operating variables. If an aircraft is maintained in a continuous full-throttle climb, at the best rate of climb airspeed, the rate of climb will be highest at sea-level; it will decrease with altitude, as engine power decreases. The aircraft will eventually arrive at an altitude where there is no excess power available for climb, then all the available power is needed to balance the drag in level flight and there will be only one airspeed at which level flight can be maintained. Below this airspeed the aircraft will stall. This altitude is the aircraft's absolute ceiling. However, unless trying for an altitude record, there is no point in attempting to climb to the absolute ceiling so the aircraft's service ceiling should appear in the aircraft's performance specification. The service ceiling is the altitude at which the rate of climb falls below 100 feet per minute; this is generally considered the minimum useful rate of climb. This diagram of forces in a climb and the subsequent mathematical expressions, have been simplified, aligning the angle of climb with the line of thrust. In fact the line of thrust will usually be 4 to 10° greater than the climb angle. The climb angle (c) is the angle between the flight path and the horizontal plane. The relationships in the triangle of forces shown are: Lift = weight × cosine c Thrust = drag + (weight × sine c) In a constant climb the forces are again in equilibrium, but now thrust + lift = drag + weight. Probably the most surprising thing about the triangle of forces in a straight climb is that lift is less than weight. For example, let's put the Jabiru into a 10° climb with weight = 4000 N. (There is an abridged trig. table at the end of this page.) Then, Lift = W cos c = 4000 × 0.985 = 3940 N It is power that provides a continuous rate of climb, but momentum may also be used to temporarily provide energy for climbing; see 'Conserving aircraft energy' below. It is evident from the above that in a steady climb, the rate of climb (and descent) is controlled with power, and the airspeed and angle of climb is controlled with the attitude and particularly the included angle of attack. This is somewhat of a simplification, as the pilot employs both power and attitude in unison to achieve a particular angle and rate of climb or descent. The angle of attack in a climb is the pitch attitude minus the angle of climb being achieved plus the wing incidence. A very important consideration, particularly when manoeuvring at low level at normal speeds, is that the steeper the climb angle the more thrust is required to counter weight. For example, if you pulled the Jabiru up into a 30° 'zoom' climb the thrust required = drag + weight × sine 30° (= 0.5) so the engine has to provide sufficient thrust to pull up half the weight plus overcome the increased drag due to the increased aoa in the climb. Clearly, this is not possible, so the airspeed will fall off very rapidly and will lead to a dangerous situation if the pilot is slow in getting the nose down to an achievable attitude. Never be tempted to indulge in zoom climbs — they are killers at low levels. 3.2.3 Forces in a descent If an aircraft is cruising at, for instance, the maximum 75% power speed and the pilot reduces the throttle to 65% power, the drag now exceeds thrust and the pilot has two options — maintain height, allowing the excess drag to slow the aircraft to the level flight speed appropriate to 65% power; or maintain the existing speed and allow the aircraft to enter a steady descent or sink. The rate of sink (a negative rate of climb, or altitude lost per minute) depends on the difference between the 75% power required for level flight at that airspeed and the 65% power utilised. This sink rate will remain constant as long as the thrust plus weight, which are together acting forward and downward, are exactly balanced by the lift plus drag, which are together acting upward and rearward. At a constant airspeed, the sink rate and the angle of descent will vary if thrust is varied. For example, if the pilot increased thrust but maintained constant airspeed, the rate of sink will decrease — even becoming positive; i.e. a rate of climb. If the pilot pushed forward on the control column to a much steeper angle of descent, while maintaining the same throttle opening, the thrust plus weight resultant vector becomes greater, the aircraft accelerates with consequent increase in thrust power and the acceleration continues until the forces are again in equilibrium. Actually, it is difficult to hold a stable aircraft in such a fixed angle 'power dive' as the aircraft will want to climb — but an unstable aircraft might want to 'tuck under'; i.e. increase the angle of dive, even past the vertical. We discuss the need for stability in the 'Stability' module. When the pilot closes the throttle completely, there is no thrust, the aircraft enters a gliding descent and the forces are then as shown in the diagram on the left. In the case of descent at a constant rate, the weight is exactly balanced by the resultant force of lift and drag. From the dashed parallelogram of forces shown, it can be seen that the tangent of the angle of glide equals drag/lift. For example, assuming a glide angle of 10° (from the abridged trigonometrical table below, the tangent of 10° is 0.176), the ratio of drag/lift in this case is then 1:5.7 (1/0.176 =5.7). Conversely, we can say that the angle of glide depends on the ratio of lift/drag [L/D]. The higher that ratio is, then the smaller the glide angle and consequently the further the aircraft will glide from a given height. For example, to calculate the optimum glide angle for an aircraft with a L/D of 12:1. Drag/lift equals 1/12, thus tangent = 0.08 and, from the trigonometrical table, the glide angle = 5°. Although there is no thrust associated with the power-off glide, the power required curve is still relevant. The minimum drag airspeed shown in that diagram is roughly the airspeed for best glide angle and the speed for minimum power is roughly the airspeed for minimum rate of sink in a glide. This is examined further in the 'Airspeed and the properties of air' module. It may be useful to know that in a glide, lift = weight × cosine glide angle and drag = weight × sine glide angle. There is further information on glide angles and airspeeds in the lift/drag ratio section of module 4. 3.2.4 Turning forces Centripetal force When an aircraft turns in any plane, an additional force must be continuously applied to overcome inertia, particularly as an aircraft's normal tendency is to continue in a straight line. This is achieved by applying a force towards the centre of the curve or arc — the centripetal force — which is the product of the aircraft mass and the acceleration required. Remember that acceleration is the rate of change of velocity — either speed or direction, or both. The acceleration, as you know from driving a car through an S curve, depends on the speed at which the vehicle is moving around the arc and the radius of the turn. Slow speed and a sweeping turn involves very little acceleration. But high speed and holding a small radius involves high acceleration, with consequent high radial g or centripetal force and difficulty in holding the turn. Even when an aircraft enters a straight climb from cruising flight, there is a short transition period between the straight and level path and the straight and climbing path, during which the aircraft must follow a curved path — a partial turn in the vertical plane. An aircraft turning at a constant rate turn is continuously accelerating towards the centre of the turn. The acceleration towards the centre of the turn is V²/r m/s². The centripetal force required to produce the turn is m × V²/r newtons, where m is the aircraft mass in kilograms and r is the turn radius in metres. Note this is aircraft mass, not weight. Turn forces and bank angle The diagram below shows the relationships between centripetal force, weight, lift and bank angle. In a level turn, the vertical component of the lift (Lvc) balances the aircraft weight and the horizontal component of lift (Lhc) provides the centripetal force. (Note: in a right-angle triangle the tangent of an angle is the ratio of the side opposite the angle to that adjacent to the angle. Thus, the tangent of the bank angle is equal to the centripetal force [cf] divided by the weight — or tan ø = cf/W. Or, it can be expressed as tan ø = V²/gr . In the diagram, I have created a parallelogram of forces so that all horizontal lines represent the centripetal force or Lhc and all vertical lines represent the weight or Lvc.) Let's look at the Jabiru, of mass 400 kg, in a 250 m radius horizontal turn at a constant speed of 97 knots or 50 m/s: Centripetal acceleration = V² / r = 50 × 50 / 250 = 10 m/s² Centripetal force required = mass × V² / r = mass × 10 = 400 × 10 = 4000 N The centripetal force of 4000 N is provided by the horizontal component of the lift force produced by the wings when banked at an angle from the horizontal. The correct bank angle depends on the airspeed and radius; think about a motorbike taking a curve in the road. During the level turn, the lift force must also have a vertical component to balance the aircraft's weight, in this case it is also 4000 N. But the total required force is not the sum of 4000 N + 4000 N = 8000 N; it is less and we have to find the one — and only one — bank angle where Lvc is equal to the weight and Lhc is equal to the required centripetal force. What then will be the correct bank angle (ø) for a balanced turn? Well, we can calculate it easily if you have access to trigonometrical tables. If you haven't then refer to the abridged version below. So, in a level turn requiring 4000 N centripetal force with weight 4000 N, the tangent of the bank angle = cf/W = 4000/4000 = 1.0, and thus (from the table) the angle = 45°. Actually, the bank angle would be 45° for any aircraft of any weight moving at 97 knots in a turn radius of 250 metres — provided the aircraft can fly at that speed, of course. (Do the sums with an aircraft of mass 2500 kg, thus weight = 25 000 N.). Now, what total lift force will the wings need to provide in a level turn if the actual weight component (aircraft plus contents) is 4000 N and the radial component also 4000 N? Resultant total lift force = actual weight divided by the cosine of the bank angle or L = W / cos ø. Weight is 4000 N, cosine of 45° is 0.707 = 4000/0.707 = 5660 N. The load on the structure in the turn is 5660/4000 = 1.41 times normal, or 1.41g. Alternatively the 'load factor' = 1/cosine (bank angle); so, cosine 45° is 0.707 = 1/0.707 = 1.41g. Manoeuvring loads In aviation usage, the lowercase 'g' denotes the acceleration caused by the force of gravity. When an aircraft is airborne maintaining a constant velocity and altitude — the total lift produced equals the aircraft's weight and that lift force is expressed as being equivalent to a '1g' load. Similarly, when the aircraft is parked on the ground, the load on the aircraft wheels (its weight) is a 1g load. Any time an aircraft's velocity is changed, there are positive or negative acceleration forces applied to the aircraft and felt by its occupants. The resultant manoeuvring load is normally expressed in terms of g load, which is the ratio of all the aerodynamic forces experienced during the acceleration to the aerodynamic forces existing at the normal 1g level flight state. You will come across terms such as '2g turn' or 'pulling 2g'. What is being implied is that during a particular manoeuvre the lift force is doubled and a radial acceleration is applied to the airframe — for the Jabiru a 2g load = 400 kg × 20 m/s² = 8000 N. The occupants will also feel they weigh twice as much. This is centripetal force and 'radial g'; it applies whether the aircraft is changing direction in the horizontal plane, the vertical plane or anything between. You may also come across mention of 'negative g'. It is conventional to describe g as positive when the lift produced is in the normal direction relative to the aircraft. When the lift direction is reversed, it is described as negative g. Reduced g and negative g can occur momentarily in turbulence. An aircraft experiencing a sustained 1g negative loading is flying in equilibrium, but upside down. It is also possible for some high-powered aerobatic aircraft to fly an 'outside' loop; i.e. the pilot's head is on the outside of the loop rather than the inside, and the aircraft (and its very uncomfortable occupants), will be experiencing various negative g values all the way around the manoeuvre. It can be a little misleading when using terms such as 2g. For instance, let's say that a lightly loaded Jabiru has a mass of 340 kg, and if you again do the preceding centripetal force calculation in a 45° banked turn using 340 kg mass you will find that the centripetal acceleration is 10 m/s², centripetal force is 3400 N, weight is 3400 N and total lift = 4800 N. The total lifting force is 15% less than in the 400 kg mass calculation but it is still a 1.41g turn; i.e. the ratio 4800/3400 = 1.41. Rather than thinking in terms of ratios, it may be appropriate to consider the actual loads being applied to the aircraft structures. The norm is to use the lift load produced by the wing as a primary structural load reference. In the 400 kg mass calculation the load produced is 5660/8 = 707 N/m², compared to the 500 N/m² load in normal cruise. However, even if the total weight of the aircraft changes, the forces experienced individual structural items — the engine mountings for example — will vary according to the g force produced by the wings. Increasing the lift force in a turn You might wonder how does the Jabiru increase the lift (or more correctly, the aerodynamic force) if it maintains the same cruise speed in the level turn? Well, the only value in the equation — lift = CL × ½rV² × S — that can then be changed is the lift coefficient. This must be increased by the pilot increasing the angle of attack. (Conversely if CL — the angle of attack — is increased during a constant speed manoeuvre the lift — and consequently the aerodynamic load factor — must increase.) Increasing aoa will also increase induced drag, so that the pilot must also increase thrust to maintain the same airspeed. Thus, the maximum rate of turn for an aircraft will also be limited by the amount of additional power available to overcome induced drag. The radius of turn = V²/g tan ø metres. For a level turn, the slowest possible speed and the steepest possible bank angle will provide both the smallest radius and the fastest rate of turn. However there are several limitations: When the steepest bank angle and slowest speed is applied the necessary centripetal force for the turn is provided by the extra aerodynamic force gained by increasing the angle of attack ( or CL ) to a very high value. Also due to the lower airspeed a larger portion of the total lift is provided by CL rather than V². Consequently the induced drag will increase substantially — requiring increased thrust power and there will be a bank angle beyond which the engine/propeller will not be able to supply sufficient thrust to maintain the required lift, and thus height in the turn. All aircraft that are not certificated under the utility or aerobatic categories are limited to bank angles not exceeding 60°. A bank angle of 75° in a level turn would induce a 3.8g load factor — the load limit for a normal category certification. Similarly a level turn bank angle of 77° would induce the 4.4g load limit for an utility category aircraft. The stall speed increases with bank angle, or more correctly with load factor, thus the lowest possible flight speed increases as bank in a level turn increases. Turns at high bank angles, near the accelerated stall speed, with maximum power applied, leaves the aircraft with nothing in reserve. Any mishandling or turbulence may result in a violent wing and nose drop with substantial loss of height. (For more information on turn physics see 'Turning back — procedure and dynamics'.) If you consider an aerobatic aircraft weighing 10 000 N and making a turn in the vertical plane —such as a loop — and imagine that the centripetal acceleration is 2g; what will be the load factor at various points of the turn? Actually, the centripetal acceleration varies all the way around because the airspeed and radius must vary. For simplicity we will ignore this and say that it is 2g all around. If the acceleration is 2g then the centripetal force must be 20 000 N all the way around. A turn in the vertical plane differs from a horizontal turn in that, at both sides of the loop, the wings do not have to provide any lift component to counter weight, only lift for the centripetal force — so the total load at those points is 20 000 N or 2g. At the top, with the aircraft inverted, the weight is directed towards the centre of the turn and provides 10 000 N of the centripetal force while the wings need to provide only 10 000 N. Thus, the total load is only 10 000 N or 1g, whereas at the bottom of a continuing turn the wings provide all the centripetal force plus counter the weight — so the load there is 30 000 N or 3g. This highlights an important point: when acceleration loads are reinforced by the acceleration of gravity, the total load can be very high. If you have difficulty in conceiving the centripetal force loading on the wings, think about it in terms of the reaction momentum, centrifugal force which, from within the aircraft, is seen as a force pushing the vehicle and its occupants to the outside of the turn and the lift (centripetal force) is counteracting it. Centrifugal force is always expressed as g multiples. Wing loading — W/S The term 'wing loading' has three connotations. The prime connotation is the standard expression — design W/S (usually just 'W/S', pronounced 'w-over-s') — which is the ratio of the aircraft designer's maximum allowable take-off weight [W] to the gross wing area . (There are some complications when national regulations specify a maximum allowable weight for an aircraft category that is lower than the design weight of a particular aircraft type; see the 'Weight and balance' module.) Aircraft with low W/S have lower stall speeds than aircraft with higher W/S — so consequently have shorter take-off and landing distances. High W/S aircraft are less affected by atmospheric turbulence. W/S is expressed in pounds per square foot [psf] or kilograms per square metre [kg/m²]. The second wing loading connotation is as the operating wing loading; if the aircraft takes off at a gross weight lower than the designer's maximum, then the operating wing loading — in level unaccelerated flight — will also be lower than the design W/S, as will its stall speed. The third is the load applied by the pilot in manoeuvring flight. As we saw above, pulling 2g in a steep turn will produce a manoeuvring wing loading that is double the operating wing loading. So, if a pilot takes off in an overloaded aircraft (i.e. the aircraft's weight exceeds the design MTOW) and conducts a 2g steep turn, then that manoeuvring wing loading will be greater than the designer's expectations. 3.2.5 Limiting loads and ultimate loads Manoeuvring loads and gust induced loads To receive type certification the design of a general or recreational aviation factory-built aircraft must conform to certain airworthiness standards among which the in-flight manoeuvring loads and the loads induced by atmospheric turbulence, that the structure must be able to withstand, are specified. The turbulence loads are called the gust-induced loads. The U.S. Federal airworthiness standards FAR Part 23 are the recognised world standards for light aircraft certification and the following are extracts [emphasis added]: Three seconds is not much time, so any inflight excursion above the ultimate load will probably result in rapid structural failure. The safety factor of 1.5 applies to fairly new aircraft in good condition; as very light aircraft age aerodynamic stresses, corrosion, hard landings and inadequate maintenance contribute to reduction of that safety factor. Airworthiness certification categories Light aeroplanes can be certificated in one or more of four airworthiness categories — 'normal', 'utility', 'acrobatic' and 'light sport aircraft' (LSA). The minimum positive limit flight load factor that an aircraft in the normal certification category (at maximum gross weight) must be designed to withstand is 3.8g positive. The LSA category minimum positive limit load is 4g. The negative limit flight load factor is –1.5g for the normal category and –2g for the LSA category. Recreational aviation aeroplanes, which are limited to banked turns not exceeding 60°, generally fit into either the normal category or the LSA category. The ultimate loads for the normal category are +5.7g and –2.25g and, for the LSA category, +6g and –4g. Amateur builders should aim to meet the same minimum values for limiting load and ultimate load factors. The 'utility' category (which includes training aircraft with spin certification) limit loads are +4.4g and –2.2g while the 'acrobatic' category (i.e. aircraft designed to perform aerobatics) limit loads are +6g and –3g. Sailplanes and powered sailplanes are generally certificated in the utility or acrobatic categories of the European Joint Airworthiness Requirements JAR-22, which is the world standard for sailplanes; aerobatic sailplanes have limit loads of +7g and -5g. The 'flight load factor' calculation is defined as the component of the aerodynamic force acting normal (i.e. at right angles) to the aircraft's longitudinal axis, divided by the aircraft weight. A positive load factor is one in which the force acts upward, with respect to the aircraft; a negative load factor acts downward. The inflight load factor is a function of wing loading, dynamic pressure and the aoa, i.e. lift coefficient, but see the flight envelope. It should not be thought that aircraft structures are significantly weaker in the negative g direction. The normal level flight load is +1g so with a +3.8g limit then an additional positive 2.8g acceleration can be applied while with a –1.5g limit an additional negative 2.5g acceleration can be applied. The manufacturer of a particular aircraft type may opt to have the aircraft certificated within more than one category, in which case there will be different maximum take-off weights and centre of gravity limitations for each operational category. See weight/cg position limitations. The sustainable load factors only relate to a new factory-built aircraft. The repairs, ageing and poor maintenance that any aircraft has been exposed to since leaving the factory may decrease the strength of individual structural members considerably. Read the current airworthiness notices issued by the RA-Aus technical manager. 3.2.6 Conserving aircraft energy Energy available An aircraft in straight and level flight has: linear momentum — m × v [kg·m/s] kinetic energy (the energy of a body due to its motion) — ½mv² [joules or newton metres (N·m)]; remembering that 'm' in the ½mv² term represents mass (Note: normally, the newton metre — the SI unit of moment of force — is not used as the measure of work or energy; however throughout this guide, it is more helpful to express the kinetic energy in the N·m form rather than joules — the N·m and the joule are dimensionally equivalent) gravitational potential energy — in this case, the product of weight in newtons and height gained in metres chemical potential energy in the form of fuel in the tanks air resistance that dissipates some kinetic energy as heat or atmospheric turbulence. To simplify the text from here on, we will refer to 'gravitational potential energy' as potential energy and 'chemical potential energy' as chemical energy. We can calculate the energy available to the Jabiru cruising: • at a height of 6500 feet (2000 m) • and (air distance flown over time)= 97 knots (50 m/s) • with mass = 400 kg, thus weight = 4000 N • fuel = 50 litres. Then: • potential energy = weight × height = 4000 × 2000 = 8 million N·m • kinetic energy = ½mv² = ½ × 400 × 50 × 50 = 500 000 N·m • momentum = mass × v = 400 × 50 = 20 000 kg·m/s • chemical energy = 50 litres @ 7.5 million joules = 375 million joules. Because it is the accumulation of the work done to raise the aircraft 6500 feet, the potential energy is 16 times the kinetic energy, and is obviously an asset that you don't want to dissipate. It is equivalent to 2% of your fuel. It is always wise to balance a shortage of potential energy with an excess of kinetic energy, and vice versa. For example, if you don't have much height then have some extra speed up your sleeve for manoeuvring or to provide extra time for action in case of engine or wind shear problems. Or if kinetic energy is low (because of flying at lower speeds than normal) make sure you have ample height or, if approaching to land, hold height for as long as possible. The only time to be 'low and slow' is when you are about to touch down. However, during take-off it is not possible to have an excess of either potential or kinetic energy; thus, take-off is the most critical phase of flight, closely followed by the go-around following an aborted landing approach. Ensure that a safe climb speed is achieved as quickly as possible after becoming airborne — or commencing a go-around — and before the climb-out is actually commenced; see take-off procedure. Kinetic energy measurement Kinetic energy is a scalar quantity equal to ½mv² joules if the aircraft is not turning. The velocity must be measured in relation to some frame of reference, and when we discuss in-flight energy management, the aircraft velocity chosen is that which is relative to the air; i.e. the true airspeed. For a landborne (or about to be landborne) aircraft we are generally concerned with either the work to be done to get the aircraft airborne or the (impact) energy involved in bringing the aircraft to a halt. So, the velocity used is that which is relative to the ground. Ground speed represents the horizontal component of that velocity, and rate of climb/sink represents the vertical component. Kinetic energy, gravitational potential energy and energy conservation are complex subjects. If you wish to go further, google the search terms 'kinetic energy' and 'reference frame'. Momentum conversion Let's look at momentum conversion. Consider the Jabiru, weighing 4000 N and cruising at 97 knots (50 m/s) and the pilot decides to reduce the cruise speed to 88 knots (45 m/s). This could be accomplished by reducing thrust — below that needed for 88 knots — allowing drag to dissipate the excess kinetic energy then increasing power for 88 knots. However, if traffic conditions allow, the excess kinetic energy can be converted to potential energy by reducing power, but only to that needed to maintain 88 knots cruise, and at the same time pulling up — thus reducing airspeed but still utilising momentum — then pushing over into level flight just before the 88 knot airspeed is acquired. How much height would be gained? Consider this: • kinetic energy at 97 knots = ½mv² = ½ × 400 × 50 × 50 = 500 000 N·m • kinetic energy at 88 knots = ½mv² = ½ × 400 × 45 × 45 = 405 000 N·m • kinetic energy available = 95 000 N·m • but potential energy [N·m] = weight × height • thus height (gained) = energy available divided by weight • = 95 000 N·m / 4000 N = 24 metres = 78 feet, or 9 feet gained per knot of speed converted. If we recalculate the preceding figures — doubling the initial (100 m/s) and final velocities (90 m/s) — the height gained will increase fourfold to 96 metres, or about 18 feet per knot. Conversely, if we halve the initial velocity to about 50 knots, the height gained per knot converted is halved, to about 4 feet. Note that as mass appears in both the kinetic energy and the weight expressions, it can be ignored; thus the figures are the same for any mass. Sometimes momentum (mass × velocity) is confused with inertia (a particular quality of mass). You will come across the expression 'low inertia / high drag' applied to some recreational light aircraft. This means that although all recreational light aircraft are low-inertia aircraft, compared to other recreational light aircraft this minimum aircraft has a relatively low inertial mass combined with a relatively high parasite drag profile; thus if the thrust is reduced or fails, the drag reduces the airspeed very rapidly. This is exacerbated if the aircraft is climbing. An aluminium tube and sailcloth aircraft at one end of the spectrum may be termed 'low momentum' or 'draggy', while an epoxy composite aircraft at the other end may be termed 'slippery'; some are very slippery indeed. The standing world speed record for an aircraft under 300 kg is 213 miles per hour; that amateur-designed and amateur-built aircraft was powered by only a 65 hp two-stroke Rotax. The handling characteristics for a low inertia/low drag aircraft differ considerably from those of a low inertia/high drag (low momentum) aircraft. Abridged trigonometrical table Relationship between an angle within a right angle triangle and the sides: Tangent of angle=opposite side/adjacent Sine of angle=opposite/hypotenuse Cosine of angle=adjacent/hypotenuse Degrees Sine Cosine Tangent Degrees Sine Cosine Tangent 1 0.017 0.999 0.017 50 0.766 0.643 1.192 5 0.087 0.996 0.087 55 0.819 0.574 1.428 10 0.173 0.985 0.176 60 0.866 0.500 1.732 15 0.259 0.966 0.268 65 0.910 0.423 2.145 20 0.342 0.939 0.364 70 0.939 0.342 2.747 30 0.500 0.866 0.577 75 0.966 0.259 3.732 40 0.643 0.766 0.839 80 0.985 0.173 5.672 45 0.707 0.707 1.000 90 1.000 0 infinity Things that are handy to know Rated power is the brake horsepower delivered at the propeller shaft of a direct drive engine, operating at maximum design rpm and best power fuel/air mixture, in standard sea-level air density conditions. (In a regulatory sense the definition is a little more complex.) An engine is only operated at its rated capacity for short periods during flight, usually during take-off and the initial climb. Rated power for small aero-engines is usually expressed as brake horsepower rather than the SI unit of kilowatts. Further discussion is provided in the 'Engine and propeller performance' module. To convert horsepower to watts multiply by 745.7; or to calculate kilowatts, multiply by 0.75. Design W/S is usually between 11 and 22 psf for GA aircraft, and 4 and 12 psf for ultralights. Gross wing area includes a notional extension of each monoplane wing up to the fuselage centreline but excludes any fairings at the wing/fuselage junction. For multi-engined aircraft, with the engines enclosed in wing nacelles, the wing area would also include the area occupied by the nacelles. Stuff you don't need to know High-performance military aircraft can achieve an aoa exceeding 45°. Aerobatic pilots — and combat pilots — use a value termed specific energy, E or energy height, He. It is the potential energy plus the kinetic energy per kg of aircraft weight; i.e. He = mgh/W + ½mv²/W As W = mg, then the equation can be re-arranged as He = h + V²/2g where h = height. What it expresses is the height that could be achieved if all kinetic energy were transferred to potential energy, but it is of little interest to recreational aviation. The thermal energy content of one litre of avgas is 30 million joules. With good engine handling by the pilot, that litre can provide 10 million joules of mechanical energy to the propeller shaft of most engines. The propeller of the Jabiru is maybe 70% efficient at cruise speed and provides 7.0 million joules, or N·m, of energy from the litre of fuel. Roughly how far will that take the Jabiru cruising at 97 knots? Easy! Drag is 540 N, so 7 000 000 / 540 = 12 965 m or 7.0 air nautical miles. We specify air nautical miles because wind will affect the distance travelled over the ground. STRICT COPYRIGHT JOHN BRANDON AND RECREATIONAL FLYING (.com)

3.1.1 Introduction The four forces When a well-trimmed aircraft is cruising (i.e. flying at a constant speed, and maintaining a constant heading and a constant altitude) in non-turbulent air, there are two sets, or couples, of basic forces acting on it. The two forces in each couple are equal and approximately opposite to each other otherwise the aircraft would not continue to fly straight and level at a constant speed; i.e. the aircraft is in a state of equilibrium where all forces balance each other out so there is no change in its motion. The couple that acts vertically is the lift, generated by the energy of the airflow past the wings and acting upward, and the weight acting downward. So, being equal and approximately opposite, the lifting force being generated must exactly match the total weight of the aircraft. The couple that acts horizontally is the thrust, generated by the engine-driven propeller, and the air resistance, caused by the friction and pressure of the airflow, or drag, trying to slow the moving aircraft. The thrust, acting forward along the flight path, exactly equals the drag. The thrust provides energy to the aircraft and the drag dissipates that same energy into the atmosphere. The forces are not all equal to each other. In fact, an aircraft in cruising flight might generate ten times more lift than thrust. When all forces are in equilibrium a moving aircraft will tend to keep moving along the same flight path at the same speed — whether it is flying straight and level, descending or climbing — until an applied force or a displacement force changes that state of motion. For instance, if the pilot opens the engine throttle fully, and maintains level flight, the thrust force is initially greater than drag and the aircraft accelerates. However, as the speed of airflow over the aircraft increases, the air resistance also increases and the aircraft will soon reach the speed — its maximum — where the forces are again balanced. Inertia, momentum and energy This property of resisting any change in motion, or continuing in the same state of motion or state of rest, is inertia. The mass of a body is a measure of its inertia; i.e. its resistance to being accelerated by an applied force increases with mass. The unit of mass we will be using is the kilogram [kg]. Some older texts refer to the 'slug', which is the unit of mass in the old British gravitational system of measurement units and equals 32.174 lb. A slug accelerates by one foot per second per second when a force of one pound-force is applied to it. Air also has mass and thus inertia, and will resist being pushed aside by the passage of an aircraft. That resistance will be felt both as drag and as pressure changes on the aircraft surfaces. A moving aircraft has momentum, which is mass × velocity and is a measure of the effort needed to stop it moving. (Momentum and inertia are not synonymous). The same aircraft also has the energy of motion — kinetic energy — which is related to mass × velocity squared. Also, because it has climbed above the Earth's surface, it has acquired additional gravitational potential energy which, in this case, is weight × height gained. Energy is discussed further in the section on conserving energy in the next module. An aircraft in flight is 'airborne' and its velocity is relative to the surrounding air, not the Earth's surface. (A ground-based observer sees the aircraft movement resulting from the sum of aircraft velocity and the ambient air velocity — horizontal motion [the wind] plus vertical motion [updrafts, downdrafts and wave action].) However, when the aircraft encounters a sudden change in the ambient air velocity — a gust — inertia comes into play and momentarily maintains the aircraft velocity relative to the Earth or, more correctly, relative to space. This momentarily changes airspeed and imparts other forces to the aircraft. (The fact that inertia over-rides the physics of aerodynamics is sometimes a cause of confusion.) A more massive (heavier) aircraft has more inertia than a less massive (lighter) one, so is more resistant to random displacement forces — wind shear and turbulence. Ultralights — whose mass is less than 750 kg — are all regarded as 'low inertia' aircraft and particularly affected by acceleration loads produced by turbulence. Freedoms of movement If the aircraft's control system allows, it can rotate about each of three axes — longitudinal, lateral and normal (or vertical) — and move forward along the longitudinal axis. An aerobatic aircraft can also move backward along the longitudinal axis — in a tail slide. Rotation about the longitudinal axis is roll, about the lateral axis is pitch — as in a ship pitching in a heavy sea — and about the normal axis is yaw — again, like much in aviation, a nautical term. Other movements can include a bodily movement along the lateral axis (sideslipping, slipping or skidding) or the normal axis (rising or sinking). Thus, an aircraft has six degrees of freedom of movement — three rotational and three translational. The three axes are relevant to the aircraft and each other (not to the horizon) so that when an aircraft is steeply banked its normal axis is closer to horizontal, rather than vertical to the Earth's surface. The axes are orthogonal (at right angles to each other) and, by convention, all are represented as passing through the aircraft's centre of gravity. When manoeuvring, an aircraft may experience any combination of the rotational and translational movements; for example, it may be rolling, pitching, yawing, slipping and sinking all at the same time. Direction of forces relative to the flight path When an aircraft is in straight and level flight lift acts vertically upward with thrust and drag acting horizontally. In fact, lift acts perpendicular to both the flight path and the lateral axis of the aircraft, drag acts parallel to the flight path, and thrust usually acts parallel to the longitudinal axis of the aircraft. So if you imagine an aircraft doing a loop, as in the diagram below, you can see that at one point, when it is going up, thrust will be acting vertically upward, drag vertically downward and lift acting horizontally towards the centre of the loop. At a point on the other side of the loop the thrust acts downward, drag upward and lift again horizontally. Weight, of course, always acts from the centre of mass of the aircraft towards the centre of mass of the Earth, so on the downside of the loop, weight and thrust are acting together and the aircraft will accelerate rapidly unless thrust is reduced. You might ask yourself this: if the aircraft is using its maximum thrust when it starts the loop, how can it climb vertically when lift no longer counters weight, and there is no extra thrust available to also counter the drag plus the weight, which are both now acting downwards? The answer is extra momentum which enables the aircraft to accomplish a fast pull-up, and usually provided by the pilot, of a lower-powered aircraft, accelerating the aircraft in a shallow dive before beginning the manoeuvre. The lift only matches the weight when the aircraft is flying straight and level. When the aircraft is in a steady descent or in a steady climb the lift is a bit less than the weight. We will explore this in the climb/descent modules but just be aware that when the line of thrust is inclined above the horizon the thrust will have a vertical component; i.e. it will provide a lifting force. When the aircraft is turning in the horizontal plane or in the vertical plane as in the loop, or anywhere in between, the lift is greater than the weight. In high-performance military aircraft it can be seven or eight times greater, because the lift provides the centripetal force to make the turn. Note: it's not always true that lift and drag act relative to the flight path. Imagine an aircraft flying straight and level, which encounters a substantial atmospheric updraught. Due to inertia the aircraft will, for the first milliseconds anyway, maintain its flight path relative to the Earth. During that time the 'effective airflow' passing by the wings will no longer be directly aligned with the flight path but will have acquired a vertical component. The lift will now act at 90° to this new 'effective airflow' rather than the actual flight path, and have a significant effect. Also, the wing itself modifies the effective airflow so just for now, until we look at aerofoils and wings, it is simpler to ignore the 'effective airflow' and other concepts and stay with the flight path. 3.1.2 Vector quantities Velocities and accelerations Vector quantities have both magnitude and direction. Velocity is a vector quantity having both a magnitude (the airspeed) and a 3-dimensional direction. A force has both a direction in which it is pushing or pulling and a magnitude (in newtons [N]), thus it is a vector quantity. Momentum, having mass and velocity, is also a vector quantity, but inertia is not. Ignoring weight and friction for now; when only one force is applied to a stationary object, the object will accelerate in the same direction as the force applied. Acceleration is the rate of change of velocity, the change being either in speed or three-dimensional direction, or both. If an aircraft accelerates in a straight line from an airspeed of 25 metres/second [m/s] to 75 m/s in 10 seconds then the average change in airspeed per second is 75 –25 / 10 = 5 m/s, thus the acceleration is 5 metres per second per second [5 m/s²]. The common usage term 'deceleration', referring to a reduction in linear speed only, is generally not used in physics as, in that science, 'acceleration' has both positive and negative connotations. Resultant forces When more than one force is being applied to an object there will be a resultant force, probably imparting an initial acceleration until all forces are again balanced at a new velocity. It is common practice to estimate resultant forces non-mathematically by drawing scaled, arrowed lines to represent each vector quantity, producing the resultant of two vector quantities in a vector parallelogram. The lengths of the lines represent the magnitude of each force and the placements indicate the application points and directions. The diagram is an exaggerated representation of an unpowered aircraft in a constant rate descent, showing that the lift/drag resultant is equal and opposite to the weight vector. In the diagram the resultant shown is the net aggregation of the aerodynamic forces generated by the wings, and it is conventional to then resolve that into its lift and drag components. We will be looking at aerodynamic forces in later modules. 3.1.3 Weight Weight as a body force There is a common, if not universal, tendency to equate the mass of a body with its weight. This is not surprising, as both are usually expressed using the same unit — kilograms [kg]. You need to appreciate that weight is a 'body force', the product of the body mass and the acceleration due to gravity. The force due to gravity — or weight — of an aircraft on the ground or in flight is expressed as W = m × g, where — somewhat confusingly — m is the symbol for mass (rather than metres) and g is a gravity constant applied to objects on, or near, the Earth's surface. That constant is not a force but an acceleration of 9.806 m/s² — also known as the acceleration of free fall. In coming calculations we will use the performance of an early version of the Australian designed Jabiru aircraft as representative of the general aviation/ultralight four cylinder, two-seat, fixed-pitch and fixed-undercarriage recreational aircraft. All forces, including weight, are measured in newtons. The maximum allowed take-off weight [MTOW], of the Jabiru of mass 430 kg sitting on the runway is m = 430 × g= 9.806 = 4216 N. The Jabiru with an 80 kg pilot on board and with half the maximum fuel load would have a mass of only 340 kg which makes some difference to performance. We shall explore this in other modules. To simplify calculations, we will load full fuel, plus a lightweight passenger, into the Jabiru giving it a loaded mass of 400 kg and use g=10 m/s², thus weight = 4000 N. Centre of gravity The position of the centre of mass or centre of gravity [cg] within the aircraft will vary according to the seating of passengers and stowage of luggage. The knowledge of the total mass of the loaded aircraft and the cg position — the weight and balance — is very important, as we will see in the 'Weight and balance' module. Weight is always presumed to act from the cg position to the centre of the Earth. We will also see, in the 'Altitude and altimeters' module, that atmospheric conditions affect aircraft performance, and subsequently the appropriate MTOW. Vector quantities are sometimes very easy to calculate; for example if a Jabiru, weighing 4000 N, is cruising straight and level, then the lift force must be 4000 N pushing vertically upward. 3.1.4 Lift The lift equation When an aircraft is cruising in straight and level flight, at low altitudes, the wings are set at a small angle — 3 to 5° — to the 'flight path' (or the 'line of flight' or the 'effective airflow' or the 'free stream airflow' or the 'relative wind', all of which mean much the same thing in cruising flight in a non-turbulent atmosphere). The net sum of the aerodynamic reaction on the wing is a resultant force directed upwards and backwards. Aerodynamicists have found it convenient to resolve that resultant force into just two components: that part acting backward along the flight path is the wing drag, and that acting perpendicular to the flight path is the lift. The amount of lift, and drag, generated by the wings depends chiefly on: (a) the angle at which the wing meet the airflow or flight path — the angle of attack (b) the shape of the wing, particularly in cross-section — the aerofoil (c) the density (i.e. mass per unit volume) of the air (d) the speed of the free stream airflow; i.e. the airspeed (e) and the wing plan-form surface area. There is a standard equation to calculate lift from the wings, which will be often referred to in these notes: (Equation #1.1) Lift [newtons] = CL × ½rV² × S The expression ½rV² (pronounced half roe vee squared) represents the dynamic pressure of the airflow in newtons per square metre [N/m²]. (Please note — if a 'Symbolic' font is not available, your browser will not display the Greek letter rho, the accepted symbol for air density, and may display r or ? instead.) The dynamic pressure expression, ½rV², is very similar to the kinetic energy expression ½mv², where m = mass. Air density is mass per unit volume; i.e. kg/m³, so the dynamic pressure of the airflow is the kinetic energy per unit volume. The values in the expression are: r (the Greek letter rho) is the density of the air, item (c), in kg/m³ V² is the airspeed, item (d), in m/s S is the wing area, item (e), in m² CL (C sub L) is a dimensionless quantity — the lift coefficient — which relates mostly to item (a), but also to item (b). In normal operations for very light aircraft, and when there are no high lift devices incorporated in the wing structure, CL usually has a value between 0.1 and 1.5. It can be regarded as the ratio of the conversion of dynamic pressure into lift, by the wing, at varying angles of attack. Angle of attack and the lift coefficient Item (a) above, the angle at which the wings meet the flight path — more properly termed the geometric angle of attack — is near 16° at minimum controllable airspeed and around 2 to 5° when cruising at low altitudes; less at higher speeds, greater at higher altitudes. We will cover the close relationship between CL, angle of attack (aoa or alpha) and airspeed in the aerofoils and wings module. The diagram shows a typical CL vs angle of attack curve for a light aircraft not equipped with flaps or high-lift devices. From it you can read the CL value for each aoa, for example at 10° the ratio for conversion of dynamic pressure to lift is about 1.0. Note that CL still has a positive value (about 0.1) even when the aoa is –1°. This is because of the higher camber in the upper half of the wing; some highly cambered wings may still have a positive CL value when the aoa is as low as –4°. A light non-aerobatic aircraft pilot would not normally utilise negative aoa because it involves operating the aircraft in a high-speed descent, but we will discuss this further in the 'Flight at excessive speed' module. Also note that the lift coefficient increases in direct relationship to the increase in angle of attack, until near 16° aoa where CL reaches its maximum and then decreases rapidly as aoa passes that critical angle. A rule of thumb for light aircraft with simple wings is that each 1° aoa change — starting from –2° and continuing to about 14° — equates to a 0.1 CL change. Also, it is not just the wings that produce lift. Parts of a well-designed fuselage — the aircraft body — can also produce lift and the vertical component of the thrust vector can supplement lift when that vector is angled upwards. We can calculate CL for the Jabiru cruising at an altitude of 6500 feet and an airspeed of 97 knots (50 m/s). The wing area is very close to 8 m²: • lift = weight = 4000 N • r = 1.0 kg/m³ (the approximate density of air at 6500 feet altitude) • V² = 50 × 50= 2500 m/s • S = 8 m² Lift = CL × ½rV² × S Dynamic pressure = ½ ×1.0 × 2500 = 1250 N/m² So, 4000 = CL × 1250 × 8, — thus — CL = 0.4. Changing the angle of attack Now what happens if the pilot decides to decrease airspeed to 88 knots (45 m/s) while maintaining the same altitude? First, the pilot decreases power to give 88 knots then adjusts control pressure to maintain the same altitude — but look at the changes in the lift equation. • lift still equals weight = 4000 N • air density still = 1.0 kg/m³ • V² changes and is now = 45 × 45 = 2025 m/s • S can't change = 8 m² Dynamic pressure = ½×1.0 × 2025 = 1012.5 N/m² So, 4000 = CL × 1012.5 × 8, — and — CL = 0.5 approximately. So, the result of decreasing airspeed, while maintaining straight and level flight, is an increase in the lift coefficient from 0.4 to 0.5. That has two possible contributors — the shape of the aerofoil and the angle of attack; items (a) and (b) above. Because the pilot can't change the aerofoil shape (unless flaps are extended, which we discuss in the 'Aerofoils and wings' module) the angle of attack must have changed. How? By the pilot adjusting control pressure to apply an aerodynamic force to the aircraft's tailplane (or some other control surface), which has the effect of rotating the aircraft just a degree or so about its lateral axis. Once the pilot has achieved the desired aoa, as indicated by the new airspeed (which will be explained in the 'Airspeed' module), the tailplane trim control is adjusted and the aircraft will then maintain that aoa. It may be appropriate to slip in another slight complication at this point. Lift, like weight, may be taken as acting through a central point — the centre of pressure [cp]. The position of the cp changes with aoa and this movement has a significant effect — it causes the nose of the aircraft to pitch up or down. So, the lift and weight are usually not in equilibrium and the rotational moment must be counteracted by aerodynamic forces produced by the horizontal stabiliser. Other tailplane surfaces also produce aerodynamic forces for trim and control, so to maintain an aircraft in straight and level flight — apart from the four forces mentioned — there will always be another force, or forces, generated by the fixed tailplane of most aeroplanes or its movable surfaces. We will look at this in the 'Stability' module. 3.1.5 Thrust Action and reaction In the Jabiru the engine supplies torque directly to the propeller shaft and the propeller converts the torque to thrust; we will amplify how this is accomplished in the 'Engine and propeller performance' module. The propeller pushes backwards a tube of air with the same diameter as itself; i.e. it adds momentum to the tube of air where momentum = mass × velocity, and is also a vector quantity. Increasing the speed also imparts kinetic energy to the air. This tube of accelerated, energised air is the slipstream. Considering Isaac Newton's third law you expect an equal and opposite reaction to the action of adding momentum. This reaction is the application of forward momentum to the propeller, which pulls the rest of the aircraft along behind if the engine/propeller installation is a 'tractor' type, or pushes it if the engine/propeller installation is a 'pusher' type. The line of thrust We need to clarify the line of thrust. This line is extended forward through the propeller shaft, which is usually aligned with the longitudinal axis of the aircraft, but not always. For instance, the engine and propeller installation in the carrier-borne Hellcat, of Second World War fame, was vertically offset so that the thrust vector was 3° down; consequently, the aircraft flew with a rather jaunty tail-down attitude. I think the reason was that the thrust line then extended back over the cg, making the aircraft more stable at the very low speeds required for deck landing. (You can read a little about the deck landing techniques of those days — my youth — in this magazine article). The relationship of the longitudinal axis with the horizontal flight path — the aircraft's attitude — varies with the speed of the aircraft. At maximum allowable airspeed, or Vne, the longitudinal axis might coincide exactly with the flight path but as speed decreases, the axis starts to angle up and could be inclined 15° to the flight path at minimum controllable airspeed. Because the line of drag is always aligned with the flight path, then the thrust vector does not directly oppose the drag vector. The diagram, slightly exaggerated for clarity, shows the relationship between angle of attack and line of thrust to the flight path, for an aircraft maintaining level flight at a very slow speed. The flight path is horizontal so the drag vector will also be horizontal; i.e. aligned with the relative airflow. The line of thrust is aligned with the longitudinal axis, so the angle between the thrust line and the horizontal flight path is the aircraft attitude — in this case, its attitude in pitch. The wing chord line is extended so that the geometric angle of attack can be seen — the angle between the chord line and the flight path. The lift and weight vectors would both be at right angles to the flight path. You might notice from the diagram that the thrust vector will have quite a substantial vertical component, so that part of the thrust is supplementing lift. Thus we have just destroyed our previous assertion that if an aircraft is flying straight and level, lift must always equal weight. In this instance, the lift is less than weight and the (very small) shortfall is provided by the vertical component of thrust. So it is more correct to say that, if an aircraft is flying straight and level, lift plus the vertical component of thrust must equal weight. We can estimate the thrust delivered by the Jabiru's propeller, cruising at a speed of 97 knots at 6500 feet: From the pilot's operating handbook we find that the engine is rated at 80 hp × 746 = 60 kilowatts [kW], and cruise power for 97 knots is 65%, or 39 kW. The fixed pitch propeller is about 70% efficient at cruising speed so the effective power from the propeller is 27 kW. One watt is the work accomplished by a force of one newton moving an object one metre in one second. The aircraft is moving at 97 knots, or 50 m/s (knots × 0.514 = m/s), so, 27 000 W divided by 50 [m/s] equals 540 N and that is the thrust being provided, which also means the drag is 540 N. Compare that to the weight and lift of 4000 N, stated in section 1.3, and you see that the lift force — silently and efficiently generated simply by the angle of the wings and the velocity of the airflow — is 7.4 times the thrust force — noisily and inefficiently generated by the engine burning expensive fuel. That is being a bit unfair because the wings are really converting much of the thrust into a lifting force. Put another way the lift to drag ratio when cruising at 97 knots is 7.4:1. We will examine lift/drag ratios in the 'Aerofoils and wings' module. The slipstream The slipstream speed of an aircraft at cruising speed might be 20% greater than the aircraft speed, during a climb it could be 50% faster, and when the aircraft is maintaining height near its minimum controllable speed, slipstream velocity might be 100% greater. Some aircraft are designed so that the slipstream over the centre section of the wings increases V and thus lift, and the combination of the vertical component of thrust plus the slipstream effect means that possibly 25% of the thrust output is contributing lift when flying in a tail-down attitude. 3.1.6 Air resistance to motion Induced drag and parasite drag Drag is the resistance of the air to an aircraft pushing through it. The resistance depends on: (a) the streamlining of the aircraft body (b) i. the excrescences attached to the airframe ii. turbulence at the junctions of structural components iii. the cooling airflow around the engine (c) the roughness of the surface skin (d) the 'wetted' area; i.e. the amount of surface exposed to the airflow (e) the density of the air (f) the speed of the airflow (g) the angle of attack. These components of drag are classified in several ways and we will look at them in the 'Aerofoils and wings' module. Part of the air resistance, the induced drag, is a consequence of item (g) the angle of attack. Induced drag is very high, maybe 70% of the total, at the high aoa of the minimum controllable airspeed, but induced drag decreases as speed increases, being possibly less than 10% of the total at full throttle speed. However the balance of the air resistance, known as parasite drag, increases as speed increases until the total air resistance equals the maximum thrust that can be produced. You can see from the diagram that parasite drag is directly proportional to dynamic pressure [½rV²] while induced drag is inversely proportional to it. Thus in normal straight and level flight, air resistance is high at both minimum and maximum airspeeds and lowest at some mid-range speed where — as resistance is at a minimum — the thrust required to maintain constant height will also be at a minimum; consequently, that is the speed — Vbr — which provides maximum range. If drag is at a minimum, then the lift/drag ratio will be at a maximum; consequently, this is very close to the best engine-off glide speed — Vbg. Air density (and thus air resistance) decreases with increasing altitude. So, the parasite drag component for a given airspeed decreases with increasing altitude while the induced drag component increases, because the wing has to fly at a greater aoa to produce the lift required. The standard expression for total aircraft drag is very similar to the lift equation: (Equation #1.2) Total drag [newtons] = CD × ½rV² × S where CD is the total drag coefficient and the ratio of total aircraft drag to dynamic pressure. CD increases as aoa increases. Things that are handy to know The generic term aircraft covers a wide range of airborne vehicles: lighter-than-air (aerostats); e.g. airships and hot-air balloons, and heavier-than-air (aerodynes). The latter includes gliders, powered parachutes, weight-shift controlled trikes or microlights, rotorcraft, helicopters, unmanned aerial vehicles and aeroplanes of all types from high performance supersonic fighters, jump jets or jumbo jets, to those like our friendly single-engine Jabiru. However, for the purpose of these notes we use the term aircraft to refer only to a class of 'general aviation' and 'sport and recreational aviation' aeroplanes with piston-engines driving propellers; relatively simple wing configurations including parawings; minimum power-off controllable airspeeds possibly as low as 20 knots; and possibly a maximum cruise airspeed, for a really high-performance aeroplane in that class, of 250 knots. These aircraft generally have one or two engines rated at 20 hp to 400 hp. The general aviation aircraft have normal wing loadings (i.e. weight/wing area) between 7 and 24 pounds per square foot (35–120 kg/m²) and maximum lift/drag ratios of 9:1 to 12:1, while ultralights have wing loadings between 4 and 12 pounds per square foot (20–60 kg/m²) and aircraft lift/drag ratios of 5:1 to 12:1. A newton [N] is the force required to give a mass of one kilogram an acceleration of one m/s². Isaac Newton's third law: If one body exerts a force on another there is an equal and opposite force — a reaction — exerted on the first body by the second. A knot is a speed of one nautical mile per hour, 0.5144 m/s or roughly 100 feet per minute. One nautical mile is the length, at the Earth's surface, of one minute of arc of a great circle and currently accepted to equal 1852 metres or 6076.115 feet. Stuff you don't need to know • Objects do not fall freely in the Earth's atmosphere. The air resistance (drag) increases as both fall velocity and air density increase until a terminal velocity is reached — where the drag force and the weight (the force due to gravity) are balanced — and the object stops accelerating. If the fall continues the object will start to slow slightly because of increasing air density at lower altitude, which increases drag. A streamlined body will have a higher terminal velocity than a non-streamlined body, of the same mass, because of the lower drag. 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