Material for landing gear

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What's better for a spring type (found on the Cessna's) landing gear. A composite lay-up or fabricated from metal? I heard the best type of material for that kind of landing gear is Metal. Composite has a tendency to "twist" under load????? If you make it (composite) thick enough....why would it twist.

So.....what is it??? Or is there really a difference
 

Dust

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Well, as always my experience is limited to a cozy, but this is it's story (along with the long EZ)


The main landing gear bow is one of the few items we cannot fabricate and must buy. It is made in a form under preasure and is full length strands of S glass. One we recieve them we must then add 9 or so layers of unidirectional cloth at a 30 degree angle to strengthen for tortional stability. we then lay up 45 layer attach points.

One company tried to make these bows and they looked just fine, the problem is that a year or two later they had spread, where the recommended manufactures have not in hundreds of planes after mmany landings, i have heard that not all of them were perfect.

Chapter 9 is the chapter that you install the main landing gear in and it is the worst chapter in building a cozy, if you build a cozy, i advise NOT reading the chapter until you get to it as it could scare you off. It is one of only a few chapter numbers i remember what were for.

It shouldn't, because hundreds or thousands have gone before and done it, but it is not simple.

That being said, i will build anything, except landing gear from e glass/s glass/carbon fiber/kevlar/boron from scratch.

enjoy the build

dust
 

Midniteoyl

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Express did away with the Fiberglass gear legs in favor of Aluminum... The Glass ones just had a tendency to crack and twist way too much...I dont have the link handy for thier supplier on this computer (Orion?), but they are real nice, substantial pieces with Internal Brake lines...
 
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orion

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Landing gear design is one of those ethereal areas that you either know or don't. I do, but just enough to do the detail engineering, not the overall specifications. Remembering what I read over the years, the design of a landing gear is not only a function of the material but also of the airplane, the landing speed, the type of gear, the tires and how much they aid in dispersing the landing energy, etc. Most of the publications and articles I read though seem to reach the same conclusion - the best material for landing gear is spring steel, although aluminum is pretty good also. All authors considered composites marginal although, with the right design characteristics and methodology, most considered graphite attractive (mainly due to the potential weight savings).

While I agreed with their reasoning and math, I was at the time flying a Grumman, which of course has one of the most robust and successful laminate gear legs I've ever seen (and certified too). So much for published articles.

I think the reason many eventually go to aluminum is for time and cost issues. Making a speciality gear such as that for the Cozy is a time consuming process involving special procedures and tools, all of which are generally beyond the homebuilder. Making aluminum gear is much simpler.

Personally I like aluminum over steel since the material tends to be a bit thicker and thus, can be gun-drilled for the brake lines. The supplier I've used is Grove Aircraft in Southern California.
 

wsimpso1

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Pazmany's book is really good at describing the parameters for all of this, get it from the EAA Book Service. I have gone through gear design in some detail too, as well as being an engineer who lives in the area of vibration isolation, which means that I live in springs.

Landing gear are all about sucking up a bunch of energy while carrying a specified multiple of the airplane's weight. You actually work out the energy in terms of static drop hieght. Now if you will always land nice and gentle, you can do just about anything you want for gear energy capacity, and in Experimental Aircraft, you are allowed to. But not me, and Van's works in steel legs too. Think about that.

Spring materials can be compared by a couple of methods: Energy stored per unit weight and energy stored per unit volume. Hardened spring steel is tops for metals by both measures. High strength aluminium alloys mean both more weight and more frontal area to store the same energy. Now you can design a 3g gear leg that weighs less than the same geometry in steel, but it won't suck up as much energy. Ostensibly, you could gun drill the steel too, but you do have to find a spring vendor that will do that for you.

I designed glass legs, and they got terrible heavy and thick and the mounting structures got too thick for my wing by the time I got to energy and less than 3g. I could revive the program and use carbon if anybody is interested.

For my money, a steel leg of a constant thickness, tapered in width is about as good as you can do for weight and compactness, with a tapered tube a close second. Both work well with clamps for mounting, the axle can be attached directly to it, with the brake line attached to the trailing edge and then covered with a teardrop fairing. You don't have to figure out how to deal with brake heat, etc.
 

StRaNgEdAyS

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Ahhh the feet the thing sits on. One of the few items I have to finish up.
I have done some peliminary drawings of what I want to hold my plane off the ground, but I want to go have a close up look at some existing retractables to get a better idea of what I'll need.
I decided to use gas cartridges to provide the suspension/spring componrnt of my legs, sitting inside a telescoping strut. I may have to alter this depending on the changes I will have to make after I look about a bit more. If I get lucky enough to find an existing set of basic mechanical retracts at the right price (and weight) I'll use those instead! :D
Basically though, I have opted for alloy upper section, secured to a pivot point on the airframe, with CrMo tube holding the wheel, telescoping up inside the alloy section.
 
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Sonja

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The main advantage of a glass landing gear is it lower weight compared to metal. I have an all-glass gear on my Pulsar, and like it better than metal. The drawback is if it is overstressed, it breaks rather than bends. If it is properly designed, it breaks somewhere in the middle, not at the top where it would cause a prop strike. Metal fatigues, so if you plan to make thousands of landings composite is better. The cases where I heard of a composite gear breaking means to me it was not properly designed for the loads in service.
 

Dust

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george, said

"Here is something I can brag about"

Brag away, you have a right to, loooooookin good

enjoy the bragin

dust
 

orion

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Just a quick note regarding an above post - Sonja mentioned that the metal gear legs would fatigue so composite is better. This however is only true if you design the correct laminate and only use a material such as graphite or Boron, both of which have excellent fatigue properties (virtually a straight and flat S-n line, only a bit less than horizontal).

The "S-n" curve (stress versus number of cycles to failure) for fiberglass is for all practical purposes exactly the same shape as that for aluminum. In blunt terminology, FIBERGLASS DOES FATIGUE (this applies to uni and to woven materials, although the fatigue properties of woven fabrics are of course somewhat worse than for uni fibers). If the glass then is coupled with a resin system that exhibits brittle properties, the fatigue characteristics become a real challenge to the point of where I would urgently recommend against such a combination's use in an aircraft application.

Fatigue is one area that the kit industry has not really addressed all that well in its design practices. Part of the reason is that there is not a whole lot of composite fatigue property information publicly available. Couple this with relatively little information out there about the fatigue life of general aviation aircraft, and you get an idea of what a future potential problem may be when a lot of our airframes start approaching several thousand accumulated flight hours.

The other reason for a lack of this specific type of analysis seems to be that many designers seem to overlook this discipline as part of their design cycles. I don't know whether this is general ignorance of the problem or just an unfamiliarity with the process. Fortunately though, many kit airplanes are somewhat overdesigned and of course, unpressurized. Also, many owners really treat their kits with extreme care and gentleness. Still, I certainly would feel better if I saw a design cycle that specifically addressed this issue in at least the critical parts of the airframe.
 
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Here is how that works, take two basically same airplanes except one made from aluminum and the other from composites. If done as it should, because the aluminum yields and just deforms before it breaks, the aluminum plane will be say 5g design with 1.5 safety factor , however because the composites breaks all at once The composite plane is designed with the safety factor of two, This in turn will put the composite airplane much lower on the "S-n" curve in the service and not much of a concern.
Of course I don’t know anything about designing airplanes I just read it in some book or something.

George
 

orion

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Unfortunately fatigue and aircraft structures don't exactly work that way so a higher safety factor will not cure the issue. Furthermore, fatigue is a cumulative effect and so, if not addressed, even a conservative design that is never flown aerobatically, may actually fail in fatigue even though in service it never comes even close to its limit loads.

Another point - composites do not "break all at once", that is an urban legend based on a fairly common misconception and erroneous information. As a matter of fact, it is often aluminum structures that fail all at once since the failures are onset through material buckling, not axial yield and failure. Actually, composites can absorb and reflect a very high amount of energy and as such, if designed right, pound for pound can be substantially safer than a structure built from aluminum.

However, will composites be lighter? In a small airplane, no. Even a graphite structure will most likely be heavier than a reinforced monocoque aluminum one. But say, in an event of a "forced" landing, the composite structure is more likely to stay intact, or at least to the point of protecting the occupants, than a typical aluminum one.

As far as the safety factors are concerned, the "1.5" number we see used today (as required by the FARs) is derived from the ratio of yield and ultimate strength values of alloy steel tubes. In other words, for most steels, the ultimate strength is generally in the ballpark of 1.5 times the yield strength of the material.

For aircraft grade aluminum, this ratio is generally much less than 1.5 and so for aluminum we design for ultimate load only. Those stresses, divided by 1.5, end up being much less than yield at the "limit" load and as such, aluminum airplanes must not yield at all at the limit loading.

For a more conservative design, with only minor weight penalties, it can be a good idea to design to ultimate loading, but consider a permanent set (material yield) as a failed condition. At this point the structure has not failed but has bent. As such, in those circumstnces, we would design the airpane to the ultimate loads but would consider the yield strength of the material as the failure point. This can actually be good practice since as I said above, most aluminum structures can fail by buckling, not yeild and breakage. Buckling has no warning and the onset of columnar buckling can be at a fraction of the material's yeld strength.

For instance, thin wall tubes in bending actually fail by crippling (or buckling) the compression side of the tube. The membrane instability failure is likely to occur at as llittle as 30% of the tube's theoretical strength.

The selection of the 2.0 factor for composite structures was an arbitrary number picked from thin air (I've never seen any actual documented justification for the number from the FAA or anyone else). It's selection was primarily based on the inability of test data to confirm a narrow range of material properties from a given set of test cupons. Looking at the history of said tests though, the variability of the cupon strengths is however more a function of the test than of the materials being tested. However this would take a long discord to explain so for now I'll leave it at that.

In short, the structural design of an airplane must be able to account for as many of the issues as possible. This is why it can be a rather tedious and at times, a time consuming venture. We have great tools today for doing most of the work but there is no substitute for real knowlege and experience. Fatigue of aircraft structures is one of those areas that is often overlooked, especially in the kit industry. Exessive design safety factors can partially overcome this shortcoming but personally, I'd like to se a design program cover as many details as possible before I'd put my rear in the pilot's seat.
 

wsimpso1

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Well put Orion.

The FAA's perspective is easily summed up as "We are having such a hard time addressing fatigue, water absorption, variability of test results, and so on, that we will just apply a design factor of safety of 2.0 for composites."

Now this is not a bad thing. It looks over-designed when new, but many resins and fibers will take on water and where available, fuel components over time, and that weakens the structure. Fatigue clues are few and far between with composites, the first clue most of us will ever get is cracked paint.

Some folks will point out that we are designing to first fiber failure, and in tensile tests the load will continue to increase after first fiber failure, in more complex structures, this is not always the case. As Orion points out, with some rather dramatic buckling modes available. If for no other reason, this is why composite structures should be sandich structures with as thick a core as is practical.

One of the biggest reasons that our small composite airplanes is heavier than if it where aluminium is that we have to design the outside skins to survive handling, construction, bug and bird strikes, rain, snow, and the inevitiable idiot at a fly-in using your bird as a writing desk. Then there is making the inside skins survive shoes, baggage, etc. All of that ends up upgauging our skins way beyond what an aluminium aircraft would need. We make up for it though. Our airfoils can be accurate laminar flow foils that maintain their shape in flight, our aero surfaces are not covered with rivet heads, and out shapes are seamless reflections of what they should be, which will always produce a faster airplane and frequently a better climb despite the weight too.

But my landing gear will still be sawn to a tapered shape out of a high strength plate, edges rounded over, bent to shape and heat treated, and all surfaces smoothed, then covered with a teardrop fairing. Because it will fit and is the lightest spring that will absorb a 9 foot drop test that the FAA specifies in Part 23. I know that I am not required to
meet that test, but I will design to do just that.

Billski
 

orion

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Hi Billski;

Just a quick question - a 9 foot drop test? Wow!

I've worked on several certified gear and to date I haven't seen any drops much over three and a half feet, or so.
 

Dust

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I know, cozy, cozy cozy, but it may be interesting to compare the weight and coast of my landing gear Bow to a comparable steel and/or aluminum.

I betcha it weighs more and costs more. As, for convience sake it is off of the plane, i will try to remember to weigh it this weekend.

To me this is interesting, although you buy the bow, the Landing gear preparation and installation is by far the most dificult chapter in the construction process

I betcha metal would be, lighter, stronger, faster.

enjoy the build

dust
 

wsimpso1

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Orion,

Showing my fallibility here Orion...

It has been over two years since I did that part of the design, but it fit in with other similar products. Perhaps it was a 9 ft/s contact velocity, which IS a lot less drop hieght. Thus humbled, I shall go back to Pazmany's book and then check my calculations for spring energy.

Your humble (and humbled) friend, Billski
 

Largeprime

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Orion
I have read in many sources that composites have no known fatigue point, and not for lack of trying. Could you point to a source that explains more about composite fatigue?

I have read that the idea of "break not bend" is a structural fact of composites. If a part is stressed to fail and any component on the composite fails, more stress transferees to the rest of the composite leading to a cascading failure. This is because the yield of the parts absorbs very little energy, unlike metals, so most of the energy is transferred to the surrounding parts, often destructively.

If this information is also incorrect could you please point me to a source to read.

Thanks
 

orion

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First, as far as pointing out a specific source of data, that's pretty difficult to do since having done this for well over twenty years, most of my data and experience comes from a wide variety of sources. The chief of these comes from organizations that I used to work for such as HITCO Aerospace Structures, General Dynamics and Boeing. Other sources inculde my own research, as well as that of others who taught me along the way.

What this leads me to say unfortunately, is that to the best of my current knowlege, there really is no simple source of data to explain what I've learned and gathered over the years, it's just a compilation of many different bits of data, design guidelines, and property information.

Outside of that though, I can make a few generic comments. First, the term "composites" is being bandied about too easily and simply, as are the terms of stress, strength, fatigue, etc. A composite is a mixture of materials however, especially in our industry, the term has come to be applied to almost any laminate structure, regardless of material, type of material weave (or lack thereof), type of resin sytem, application, or structural makeup. And therein lies the problem.

Some composite laminates (meaning materials) do behave as you indicated - they have an inherent resistance to fatigue and as such, the S-n curves are nearly flat lines so that infinite life criteria (500,000,000 cycles) for tension-tension applications correspond to nearly 90% of the material's ultimate tensile strength. When they fail however, it is through crack initiation and catastrophic propagation, that is initiated with little or no warning.

But here is another significant issue to discuss - what is the loading type. Is it tension-tension, is it completely reversed (tension-compression), or is it in bending? Each of these conditions will slightly alter the long term behavior of the laminate and so, it is relatively important to understand the behavior of said structures before making an off the cuff statement about the life expectancy about any such assembly. Unfotunately, this is a subject that could take up days, if not weeks of lecture so I can't really delve into it in any but the most basic terms.

As far as the issue of energy absorption is concerned, it is important to differentiate between locallized effects such as that we discuss when talking about material test coupons, and structural assemblies as when discussing entire airframes. Two very different subjects with unfortunately similar terminology.

It is also important to look at and differentiate between the issues of loading to breakage, and loading sufficiently to fatigue (the latter particularly applies to fiberglass structures). One also needs to examine the cumulative effects of loading the said structure over the thousands of hours of service we expecet from our aircraft.

But, if I was to put forth a general statement, in a basic sense, what you say is correct and since there is no yield point for composites, we must design to ultimate conditions, with sufficient understanding of the material at hand, or at least sufficient application of the design safety factors. However, this only addresses the ultimate strength of the airframe. This must be coupled with the fatigue data in order to arrive at the most complete understanding as to what the structure will behave like ten or more years down the line.

In looking at many kit or amateur built (plans built) aircraft (especially the older ones) over the years, there is nothing scarier to me than seeing a series of cracks in the paint or Gelcoat, and having the owner explain them away as a defect in the surface coat - something that is to be sanded, refilled, and painted over. Generally this person does not believe in the issue of fatigue since according to him, he never operates the airpalne in an aerobatic or highly loaded way, and thus he concludes that this is not an issue that could possibly affect his airpalne. In my understanding though, this is an accident waiting to happen.

OK, reading the above it looks like I'm rambling a bit so sorry for the running off at the mouth. It's just difficult to cover what is an important issue but way beyond the scope of a discussion board (besides, I'm not a fan of typing).
 
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