Confused

[Deskpilot]

I've spent this afternoon, fruitlessly trying to find info on the required load tests to be applied to an experimental aircraft. Unfortunately, I got side tracked and tangled up in the rules and regs set out on the RA-Aus site. How utterly confusing. Experimental, privately built, amateur built, reference to so many CAO's, sections and sub sections. One wonders if these rules were written by lawyers or monkeys.

 

All I want to know is, if I design a single seat plane that I'd consider experimental, do I have to load test the wings, and if so, how. Is it legal/advisable to fly a wing that has been load tested. Some say no way but not all of us can afford to build 2 or 3 times. Obviously one can't test to breaking point but there must be a way of testing to say 75% of ultimate strength and the wing still be safe to use. How would one set what is the 75% number? I'm thinking in terms of a plywood 'I' beam spar with a mixture of ply and foam ribs.

 

Just an idea, I weigh 100kgs so it the spar were rested on its tips and a load of 400kgs spread alone the spar, would that equate to 4g? I've heard that one should not support by the tips but I don't understand why not.

 

Something else just struck me, should one test a single wing (one side only) to say 4g's or the pair together.

 

 

[fly_tornado]

at a guess, you take the maximum take off weight and multiply that by the max load factor + safety margin. spread proportionally along the wing by the lift factor. you need to test the wing as unit.

 

eg 600Kgs x (4g + 1g) = 3000kgs

 

thats a lot of sand bags

 

 

[Deskpilot]

eg 600Kgs x (4g + 1g) = 3000kgs

thats a lot of sand bags

Sure is, but I'm so sure you've got it right. There's no way this little budgies gunna lift that lot. My back's more important than flying, and that's saying something.

 

BTW how do you know the final AUW, guessimate?

 

 

[icebob]

Hi Deskpilot,

 

Try going to www.silentflight.com, there is a calculator there that will give you a close weight for the load testing. There also was an article on the FAA web site for wing load testing.

 

Bob.

 

 

[Guest davidh10]

Deskpilot; The reason for not suspending the wing by its tips is simply because the lift is distributed along the wing, not acting at just the tips.

 

 

[facthunter]

Wood is hard to get strength values with. Make test sample structures and get actual performance figures from them. They should be made of the same material exactly as the complete structure is intended to be built. You don't have to build an entire wing, but if you did, and load it up with sandbags, they should be distributed as the actual lift forces would occurr. Of course you put the load on the bottom of the wing, by this method, and that is not what actually happens, but the spars wouldn't know the difference. You don't have to test to destruction either. If the structure is strong enough, then you haven't damaged it and it can be used. as long as NO part has deformed/failed. The other wing being a mirror image of the first, should be as strong, provided your material is the same and the glue applied in a quality controlled fashion. You should make glue joints and test them to destruction, to make sure your technique is good. Bill Whitney used to have a disc and notes covering all this, in a clear to understand form. Try to get one. It really is good stuff. Nev

 

 

[Deskpilot]

I've got Bills DVD's. Maybe time to view them again.

 

 

[Guest Maj Millard]

Deskpilot, If you are planning to fly in this aircraft you do need to load test the wings and their attach fittings. Last one I did I used the figures for the 'utility' catorgory (Cessna, Piper etc).

 

I don't remember the actual figures but somewhere in the order of 2.5 times the suspended weight.... Suspended weight being Fuselage, fuel, pilot, engine and prop and anything else in the fuselage for flight. The weight of the wings themselves are not included, as they are supporting themselves, but you do add the weight of any fuel therein.

 

IE:(for example) suspended weight is 190 kgs X 2.5 =475 Kgs...or higher if you decide on a higher load factor.

 

The best method of testing is of the complete wing assembly attached to the fuselage, as ready for flight.

 

To only test one wing is hoping at best.

 

If the wing is fabric covered the wing can be tested uncovered, so that strain on internal (especially wood) fittings can be observed for distorsion etc.

 

If the wing is metal , or plywood covered IE: stressed skin design, than it must be completed with those skins.

 

Testing is done with the fuselage and wings inverted, and supported on the fuselage only, at suitable strong points. The wing is set at flying angle IE: with approx 5 degs of angle of attack.

 

Once the aircraft is set up inverted, and before loading the wings, strings with weights, (like small plumbobs) can be hung from the ceiling next to selected points, along the wing leading and trailing edges to show how far the wing bends under max load. About four lines each wing should be fine (front and rear spars). These stringlines should be close to, but not touching the wing.

 

All control cables or rods should be in place, in the wing, so you can check that they don't contact anything under load.

 

OK your ready. If your wing is flat bottomed you may place plywood between front and rear spars to support sand bags or whatever. Do include the weight of the plywood in your total weight.

 

Place weights mostly on the wing roots tapering outward toward the tips (as the wing is loaded in flight) Weight should be centered along the main point of load IE: over the wing spar which carries most of the load in flight. Load both wings evenly, one side to the other. Remember also that if the wing has any degree of dihedral, plus any additional bending of the wing (now going downwards at the tips) your weight may have a tendency to slide off. Make sure your weight is not allowed to slide off!.

 

Once the wing is fully loaded, leave for 24 hours. After 24 hrs you can now check your stringlines, and note wing bending, and inspect all attach points for any twisting, or signs of stress or overloading.

 

At this point I go back and pick up the tail carefully, and give it a light shake. Things will groan and creak a bit, but it is then you will know weather you want to fly the aircraft or not !.It is a lot of work, but now you can be sure that you have adequety tested your wing design correctly, and that it is strong enough to carry flight loads to the factor that you have tested................................................Maj...

 

 

[Deskpilot]

Thanks for a very complete description/instruction Maj. It seems that there are 2 trains of thought on this subject. This one, and one that states that no tested wing should ever be flown, ie scrap it and rebuild. Obviously, if any permanent distortion should arise from the testing, then yes, it's scrapped. But other wise.......? I haven't read any where before that states leave for 24 hours, fully loaded. Where does this come from? I have also read that if one wing passes the test, no need to test the other. Below is a video of a PIK-26 wing assembly being tested. Now I realize that the wing to airframe connection also needs to be tested and can therefore see the sense in what you have said.

 

The only part of my original question not answered is, how does one arrive at the weight of an (one off) experimental plane?

 

 

 

[djpacro]

... if any permanent distortion should arise from the testing ...

A wooden structure really doesn't suffer any permanent deformation (fittings bearing in the wood and crushing it locally may manifest itself in measurable deformation somewhere). Otherwise you can load wood up to nearly failure and back down again and not notice anything. Take it a tad further and it will fail.

... how does one arrive at the weight of an (one off) experimental plane ...

I'm like Ted Baillieu - I've answered this before.

[facthunter]

Is the whole plane wood?. If you were using a steel tube truss fuselage, you can calculate the material required, but you have to analyse the stresses in the whole frame and make EACH piece strong enough to take the load. Done properly, this would be adequate off the plan.

 

I thought utility was 3.8G ultimate, but there is no point guessing it. Check it out. You can go stronger if you want. If you use a plywood strip with glued wood caps and cantilever design, it can be, quite light in construction at the tips. Your greatest sheer and bending loads are at the wing root. How are you going to stop twisting? This can be done with 2 struts, or use a torsion box spar or plywood capping .

 

Re the balance, you have to add all the weights of various "bits" of your plane and do a weight and balance on each bit/section. This involves determining the CofG of each bit and its total weight and distance from a datum point. Then you can put the wings in the right place so the plane will balance without redesign or adding weights, unnecessarily. You even have to get the CofG of the wing and take it into account Nev

 

 

[Guest Maj Millard]

The whole aircraft wing support system needs to be tested as a whole, as it will be flown. Anything less and your leaving critical areas basically untested, and then relying on those areas to be up to the job during test flights. Not the best way to go if they should fail.

 

As I described it you are not testing the wing anywhere near destruction (ultimate yield). Even in the utilitary catogory (Cessna 172 etc etc) you've probabily tested 100 percent above the loads you'll encounter on the average life of the aircraft.

 

If the design is sound, and no major problems are evident during the wing loading exercise, you now have a structure proof-tested to that catogory. I would have no hesitation in flying with that tested structure, and have done with confidence in the past.

 

The 24 hour load requirement probabily comes from the US FAA structural testing guidelines. I would imagine it is a requirement to keep constant strain on wood glue joints, or in-wing brace cables, to allow for some stretching of those cables. One requirement is to pluck or test for even tensions at the end of the test period.

 

To arrive at the weight of an experimental aircraft, you need to get out the scales and calculator and start weighing things, including the pilot, to achieve your approximate aircraft take-off weight.............................................................Maj...

 

 

[RickH]

Hey Guys

 

I agree prety much with everything Maj has said I would like however to ask one question and make one point.

 

Question. if we assume that wingloading ie x lbs/sqft spread over the entire wing, why then would you taperoff the weight toward the wing tips. After all as facthunter pointed out the greatest shear and bending moments are at the wing root, however this does not translate to greatest weight being concentrated toward the root of the wing?

 

Now to digress a little the point I would like to make, when it comes to the actual test is that if you really want to know what is happening to the structure during testing then you realy must be able to measure the total deflection under load (inorder to calculate stress and hence strain) and the absolute safest way to do this is to apply strain gauges to the top and bottom surfaces of the spar. You may think that this is getting a little technical but better to be technical and alive than technically dead.

 

Regards RickH

 

 

[Guest davidh10]

...Question. if we assume that wingloading ie x lbs/sqft spread over the entire wing, why then would you taperoff the weight toward the wing tips. After all as facthunter pointed out the greatest shear and bending moments are at the wing root, however this does not translate to greatest weight being concentrated toward the root of the wing?

The weights along the wing would have to be distributed in proportion to the lift at that point in the wing. The objective is to mimic the actual stress distribution that would occur in flight. Shear and bending moments are a separate consideration IMHO.

 

 

[Guest rocketdriver]

Hey GuysQuestion. if we assume that wingloading ie x lbs/sqft spread over the entire wing, why then would you taperoff the weight toward the wing tips. After all as facthunter pointed out the greatest shear and bending moments are at the wing root, however this does not translate to greatest weight being concentrated toward the root of the wing?

Regards RickH

I read somewhere that a straight tapered wing with washout gives a spanwise lift distribution that approximates to the ideal (which is eliptical - hence the Spitfire and Heinkel 100 wing shape). This distribution of lift provides very little lift at the tip, and so minimises induced drag(which is largely caused by air flowing around the wing tip from the bottom surface (high pressure) to the top (low pressure). If there is no lift being generated, there is no pressure difference and so no vortex production ....in theory! Moving on, if the lift distribution is approximately elliptical along the span, so should be the test weighting ....

cheers

 

RD

 

 

[Guest Maj Millard]

Rick H, The actual pressure distribution along the wing is less at the tip as the airflow is at an angle toward the tip and the air departs over the tip and the Trailing edge of the wing, (positive vs negative pressure) so actual load carried is less out there. This has alrready been demonstated by pressure testing actual wings, by greater minds than I. hope that answers your question .......................................................................................................maj...

 

 

[RickH]

Hey Maj I understand what you are saying and agree with what you say re wing tip losses but if we look at airfoil theory ie infinitely long wing and constant chord and thickness then the lift distribution and hence resulting loads are distributed evenly. or am I missing something.

 

RickH

 

 

[RickH]

PS. Wing tip losses being applicable to only the outer 5-10%

 

RickH

 

 

[Guest Maj Millard]

Actual in-flight load distribution varies with each particular wing. Each wing has a different pressure distribution profile, throughout it's span. Any good theory book on airfoils will explain the principal...........................Maj...

 

 

[djpacro]

The LAA (formerly PFA) in the UK has this useful document on the very simple Schrenk method commonly used to calculate the spanwise laod distribution for small airplanes. Only 17 pages but 7 Mb pdf at http://www.lightaircraftassociation.co.uk/2010/Engineering/Design/schrenk%20approximation.pdf

 

Go to their homepage and you'll find lots of useful stuff there, for example, loadings for structural tests - an example: http://www.lightaircraftassociation.co.uk/engineering/TechnicalLeaflets/Building,%20Buying%20or%20Importing/TL%201.15%20Example%20Aircraft%20Loading%20Calculations.pdf

 

 

[turboplanner]

Right IG, it's for RC Aircraft.

 

For those who don't fly them the power top weight ratio is roughly the equivalent of a 300 hp Jab, and most of the ones I've seen flying are pulled along by thrust.

 

You would have to use very flimsy materials indeed to get wing failures we would see in a full size aircraft.

 

Maybe the principles of the calcs are the same but beware the minor differences and coefficients in any equation used in the formulae.