Dafydd Llewellyn

Nowadays, you have to be more specific as to what you mean by "rubber"; there are at least a dozen rubber-like substances in use for elastomeric components, with totally different chemical basis. I assume you mean "natural rubber" (as from rubber trees); however that has mostly been replaced by nitrile rubber, except in rubber bands and speargun rubbers and shock cord. Other common forms include Neoprene, polyether, silicone, polyurethane, viton, etc. Each of these works with some fluids but not with others. I'm no expert on this, but it's a minefield unless you know what you're doing.

Well, thanks for that - that's it in a nutshell, isn't it? The point about wind effect is a good one; however not really a problem in a glider, because of its dive brakes. Needs to be taken into account in anything that does not have an equivalent means of shedding the excess height.

See CASR Part 21.190

Good question. If it's in the TCDS, then those are the limits. How they got there, you'd have to ask the FAA.

Thanks, DJ - about time somebody answered the original question. There's a subtlety to it, though; let me see if I can explain: Rudder lock occurs when the rudder hinge moment reverses, which can happen at sufficiently large yaw angles - we define control surface hinge moments in two ways: Firstly, the tendency of the control surface to "blow back to the neutral position" that one might measure in a wind tunnel, where the aircraft is always pointing directly into wind. Secondly, the tendency of the control surface to blow hard over against its stops, if the aircraft is sufficiently yawed. If you apply, say, left rudder, keeping the wings level, the aircraft will yaw to the left, and a rudder that has no aerodynamic balance will tend to be driven to the left. Where the rudder actually ends up, if the pilot takes his feet off the pedals, is called the "free-floating angle", and it's the consequence of the balance between these two effects. If the second effect predominates, the rudder will want to go hard over, so the aircraft will remain in the yawed position, unless the pilot forces the rudder back to its central position. The second effect actually acts to reduce the effort needed to yaw the aircraft. This is what is called "rudder lock"; and it's not actually a consequence of aerodynamic balance of the rudder, but rather due to the lack of it. Suitable forms of aerodynamic balance can increase the first effect and reduce the second effect, to the point where the free floating tendency of the control surface acts to increase the resistance to yaw. You will often see rudders that have a horn balance; it may come as a surprise to discover that they take more effort on the rudder pedals to make the aircraft yaw, than does a simple rudder without any balance. A lot of small aircraft have poor control harmonisation - mainly in the form of too-light rudder forces - simply because they lack any form of aerodynamic balance on their rudders. This may possibly contribute to stall-spin accidents, because the rudder is too light for the pilot to realise that he has too much rudder applied. The design of aerodynamic balanced control surfaces is not simple, and there's an element of "black art" in it; so this is NOT a good area for an amateur designer to play around. However, the most common "fix" for rudder lock is a dorsal fin, especially one with a sharp edge along its top; any time you see one of these, it's a fair bet the aircraft either had, or the designer anticipated, rudder lock.

Well, I suppose somebody has to ask it . . . Why did it fail? Because this thread is a waste of electrons until that information is available.

Vc and Vne are STRUCTURAL LIMITATIONS. The only effect added drag has on them is the power (or dive angle) necessary to reach them.

Yes, the C177 was certificated to a very early version of FAR 23 - and I agree there's little practical difference between the two; the gust load formulae differ somewhat, and so does the nominal design gust velocity - but the resultant wing & tail loads are about the same. The effect of having a single Va based on the MTOW is that you must design things like the engine mount for the load factors it gets at minimum weight. I do not use the methods of FAR 23 Appendix A & B to calculate the basic loads; those formulae constrain the configuration and the wing aspect ratio too much. When one comes to calculating the basic loads from first principles, it is necessary to calculate the balance and checked manoeuvre loads, as well as the gust loads, for all the corners of the flight envelope, for maximum and minimum weight, and forward and aft CG. So there are hundreds of load cases in the spreadsheet, and the critical cases for the engine mount (for example) are not necessarily the same as those for the main lift truss or the tailplane. OR you can design the engine mount for the case that happens to be critical for the wings, and step-down the value of Va for lighter weights, so that (assuming the pilot observes this - which is quite unrealistic) in theory the same design load will suffice for all weights. Similarly, you can work out the gust loads for Vc at the maximum design altitude, and have a single value of Vc - OR you can step-down Vc at increased altitude, to keep the gust loads from exceeding the sea-level values. Whenever I see either a range of values of Va according to weight, or a range of values of Vc according to altitude, I know that I'm looking at a structurally pared-to-the bone aircraft. A design that does not do these things has a little more "hidden safety margin" - and a lot of pilots place unthinking reliance on that. This thread ignores the fact that, at the limit load, the structural strength of a new airframe must not suffer any permanent deformation. It must be able to carry 1.5 times that (for a metal aircraft) without breaking. A composite structure also has a "material variability" factor on top of that. These factors confer the appearance of robustness (provided you do not have a minimum-tolerance specimen in the critical part of the structure) that some pilots abuse. It is stupid to do so, because even if the aircraft you happen to be flying is above the minimum tolerance, applying unnecessary stresses to it cause accelerated fatigue damage. A ten percent increase in stress level roughly corresponds to a halving of the fatigue life. The fatigue life of most recreational aircraft is not required to be specified (s0 far); do you REALLY want to find out how short it can be?

For a 4 - G limit-load aeroplane, twice the flaps-up stall speed is a pretty realistic approximation.

Here's the corresponding extract from CAR 3 (the design standard for the Cessna 172, 177 etc): § 3.184 Design air speeds. The design air speeds shall be chosen by the designer except that they shall not be less than the following values: Vc (design cruising speed) = 38 Ö W/S (NU) = 42 Ö W/S (A) except that for values of W/S greater than 20, the above numerical multiplying factors shall be decreased linearly with W/S to a value of 33 at W/S=100: And further provided, That the required minimum value need be no greater than 0.9 Vh actually obtained at sea level. Vd (design dive speed) =1.40 Vc min (N) =1.50 Vc min (U) =1.55 Vc min (A) except that for values of W/S greater than 20, the above numerical multiplying factors shall be decreased linearly with W/S to a value of 1.35 at W/S=100. (Vc min is the required minimum value of design cruising speed specified above.) Vp (design maneuvering speed) = Vs (Sqrt) n (my insertion - the software did not convert the square root sign) where: Vs =a computed stalling speed with flaps fully retracted at the design weight, normally based on the maximum airplane normal force coefficient, CNA. N= limit maneuvering load factor used in design, except that the value of Vp need not exceed the value of Vc used in design. Therefore, there is scope for the designer to declare different values of design manoeuvring speed at different weights - but this is normally not done nowadays in my experience. It complicates the flight envelope vastly with negligible practical benefit. The flight envelopes specified by some "simplified" recreational design standards do not require gust loads to be separately considered; therefore for those aircraft Vc and Va have to be essentially the same. The green band (where marked) on the ASI normally goes to Vc; but some design standards do not require ASI colour-coding. Aircraft such as the Sabre and the Foxbat (and the early Cessna 210s - I don't know about the later ones) are quite capable of cruise speeds significantly above Vc, which in Australian summer conditions is downright dangerous.

Va is always calculated for the MTOW - so only one value.

From FAR 23.335: © Design maneuvering speed VA. For VA, the following applies: (1) VA may not be less than VS√ n where— (i) VS is a computed stalling speed with flaps retracted at the design weight, normally based on the maximum airplane normal force coefficients, CNA; and (ii) n is the limit maneuvering load factor used in design (2) The value of VA need not exceed the value of VC used in design. (Note: Vc is the speed at which the aircraft is designed to meet the 50 ft/sec gust requirements. Some "simplified" recreational aircraft standards do not use it. and Horizontal Stabilizing and Balancing Surfaces § 23.421 Balancing loads. (a) A horizontal surface balancing load is a load necessary to maintain equilibrium in any specified flight condition with no pitching acceleration. (b) Horizontal balancing surfaces must be designed for the balancing loads occurring at any point on the limit maneuvering envelope and in the flap conditions specified in §23.345. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13089, Aug. 13, 1969; Amdt. 23–42, 56 FR 352, Jan. 3, 1991] § 23.423 Maneuvering loads. Each horizontal surface and its supporting structure, and the main wing of a canard or tandem wing configuration, if that surface has pitch control, must be designed for the maneuvering loads imposed by the following conditions: (a) A sudden movement of the pitching control, at the speed VA, to the maximum aft movement, and the maximum forward movement, as limited by the control stops, or pilot effort, whichever is critical. (b) A sudden aft movement of the pitching control at speeds above VA, followed by a forward movement of the pitching control resulting in the following combinations of normal and angular acceleration: ------------------------------------------------------------------------ Normal Condition acceleration Angular acceleration (n) (radian/sec2) ------------------------------------------------------------------------ Nose-up pitching............... 1.0 +39nm÷Vx(nm-1.5) Nose-down pitching............. nm -39nm÷Vx(nm-1.5) ------------------------------------------------------------------------ So Va is a number of things: Firstly, it's the highest speed at which the wing can be pulled to stalling incidence, in longitudinal equilibrium (i.e. by steady application of the controls) without exceeding the limit load. Note that the wings will be carrying n times the aircraft weight PLUS whatever download is necessary at the tailplane to hold that flight condition. Secondly, it's the highest speed at which any ONE control surface may be abruptly deflected to its stops (momentarily) without overloading the control surface or its support structure or the control system. However, it is still possible to overload the airframe at Va if the aircraft encounters a gust load at the same time as the manoeuvre load; and also by deflecting more than a single set of controls simultaneously. Further, the aircraft has to be strong enough to withstand this load only ONCE in its life.

Slightly off topic, but a historical note - when I was doing my basic training, Moorabbin was an "all over" grass field; do you remember the procedures for that kind of airfield? You simply turned base after the fellow in front, and then turned final so your landing run would be 50 yards to the right of the aircraft ahead of you; on landing, you taxied straight ahead to the perimeter track, and taxied back around the perimeter - which was one-way traffic, in the circuit direction. If the fellow in front was already at the far edge of the field, you started afresh at the near edge, because by then the aircraft on that bit of the field would have reached the perimeter track. I've been No. 12 on final, at Moorabbin. 5 or 6 on final was fairly typical. It didn't cause any problems. Moorabbin was at that time the busiest airfield (in terms of movements per hour) in the southern hemisphere. However, several years later Moorabbin got a sealed runway, so those procedures were dropped and it was no longer the busiest airfield in the southern hemisphere. Why? Because the airport manager got promoted a grade in the public service, if he was in charge of a sealed runway.

G Good point.

Higher than at full power? The dive brakes on gliders are very effective, because their area is large in relation to the aircraft weight ( the effective area is that of the top and bottom surface bakes plus the frontal area of the wing in between them), but small in relation to the wing area - so they give a lot of braking for relatively little loss of lift. People often ask why light aircraft don't have them; it's because the loss of lift is higher, and the braking effect much smaller, in relation to the aircraft weight.

Yes; however the equivalent in a powered aircraft is side-slipping, which gliders don't do all that well. Or if you have a constant-speed prop., shut the throttle in fine pitch. These are tools one needs to know how to use. I particularly dislike powered aircraft in which side-slipping (or "forward" slipping, if you want to be pedantic) is prohibited with the flaps extended.

Those who don't study history . . . The 1960s boom period of light aircraft production in the U.S.A. came to a grinding halt in the late 1970 / early '80s because of the U.S. Product Liability laws, and the abuse of them by the "litigious society". The "Big Three" stopped building small, single-engine light aircraft, due to the ridiculous cost of product liability insurance. So we have been soldiering on for forty years, patching up aircraft that were originally designed for about a ten year useful life. As a result, these aircraft nowadays regularly incur repair costs that exceed the original purchase price. They are way past their "use by" date. The high maintenance costs are not the fault of the GA maintenance industry; it's the consequence of trying to keep operating "planned obsolescence" products well past their design lifetimes, as a consequence of the greed of the litigious society. The maintenance industry has actually worked wonders in keeping the damn things flying at all. I owned a PA 28-140 in the '80s - purchased it third-hand for $10,500 in 1980 - and I was able to claim the operating costs against my income tax. I sold it in 1992, for $ 16,000, because I could no longer use it as a business tool. Not counting engine/propeller overhaul costs - the Lyc 0-320 cost around $15K back then, to overhaul - it cost, on average, $4000 per year plus fuel. You could factor those costs up to to-day's prices, but that would still under-estimate the costs, because by now that airframe would be rotten with corrosion. Also, the like of Lycoming etc engines can be overhauled once (if you do it when it's due) and remain reliable, and twice - if you look after them really well, like fly at least once a week - but the crankcase alloy gradually loses strength, and by the end of the third life, it's history. So engines that have been maintained "on condition" are a very poor purchase proposition. That's the situation with most GA aircraft of the sort you are talking about, nowadays. To compound that, those aircraft were not required to have any declared fatigue life, when they were certificated. That omission by the regulatory authorities has had two results: Firstly, an unrealistic expectation that they would last indefinitely; and secondly, a belated "catch-up" by the regulatory authorities via Airworthiness Directive action. Secondly, the introduction of SIDs by some manufacturers - which are really a campaign to get these old aircraft scrapped, in disguise, because the difficulties for manufacturers in supplying spares for products they stopped building forty years ago are immense. Manufacturers who have not introduced SIDs will have their products scrapped by the "ageing aircraft" campaigns currently being run by most National Airworthiness Authorities. So, there is no such thing as a "cheap" GA aircraft. The odd one that has been lovingly restored will cost accordingly. If you want to find one to restore, a tube-and-rag type may sound the least-worst option - but CASA has put the cost of maintaining an aircraft welder's licence through the roof, so that avenue isn't simple, either.

And that's something to complain about?

Mil-H-5606 stays liquid down to around minus 50 C - which means aircraft hydraulic systems remain functional at high altitude (seems like a good idea, don't you think?). Since undercarriages tend to take a while to warm up after cruising at altitude, I suppose it seemed like a good idea for them, too. As a fringe benefit, you don't tend to get the sort of local pitting corrosion that plagues automotive brakes. I think automotive systems started using the glycol/alchohol type brake fluid (to replace castor oil) because it allowed the continued use of natural rubber brake cylinder cups - whereas aircraft went to synthetic rubber (Buna-N - Neoprene) about the time they stopped using castor oil as hydraulic fluid, in the 1940s. The practice has continued to this day - tho synthetic rubber is now cheaper than natural rubber - in fact, I expect brake cups are nowadays mostly made from nitrile rubber. However if you want the metal parts of the brake system to last indefinitely, mineral-base brake fluid is definitely a step in the right direction. Some car manufacturers have seen the light - Citroen for one; they changed to LHM (liquide huile minerale) long ago - about the time they introduced the DS series, in fact. Their brake calipers just don't give trouble, in my experience. However, most car manufacturers are interested in planned obsolescence, so brake parts that corrode out are probably viewed favourably. I use "regular" brake fluid as a mild paint stripper, to remove the decorative paint trim from my aircraft without damaging the polyurethane anti-corrosion finish; it's great for that purpose. Not much good for anything else, IMHO.

Quite a bit more than that, actually, tho the debate on EFATO seems to show that as a major failing in pilots without gliding experience.

Maule M5-180

Most aircraft hydraulic braking systems are designed for mineral oil. This is NOT compatible with automotive brake fluid. The norm is MIL-spec hydraulic fluid MIL-H-5606; any aircraft maintenance organisation will have some. Auto transmission fluid is a reasonable substitute if you do not operate in extremely cold conditions.

Yes, the blue-head 582 does indeed have a bypass-type thermostat - and if the size of the bypass line is anything to go by, it is designed to handle the full water-pump flow. So when the thermostat starts to open, the cold water that starts to flow from the radiator is fed into the hot recirculating flow, and the mixture of the two goes into the base of the cylinder block. Sounds to me like as good a fix for shock cooling as one is likely to find. And you get the benefit of a fast warm-up.

Most current aircraft design standards require that, without exceeding the aircraft's certificated MTOW, you must be able to fill all the tanks, with only the pilot (a standard weight pilot, usually 86 Kg nowadays); AND you must be able to carry fuel for one hour at normal cruise power, with all the seats full (to their placarded or stated occupant weight limits). If the tankage were limited so that you could fill all the tanks and all the seats, the available range at reduced load would be unnecessarily restricted.

Yes.