Essentially. The stress report describes this in the appendix. I don't think a test with covered wings is a good idea however. I'd want to observe the structure as the loading progresses.
Bob, I have seen some pictures of load testing of cantilevered metal wings, where the load ( of sand bags ) was applied on top of the wing. The Max stress report states load should be applied on the underside of the wing ( with aircraft in inverted position ).
Is this change necessitated because because Max is a strutted airplane, and only in the inverted position, the struts will carry loads ( in tension ?) like in normal flight ?
I am interested in the subject of load testing. It really is unbelievable the amount of load a wing can carry, when you see so many sand bags piled up atop the wing.
Bob, I have seen some pictures of load testing of cantilevered metal wings, where the load ( of sand bags ) was applied on top of the wing. The Max stress report states load should be applied on the underside of the wing ( with aircraft in inverted position ).
Is this change necessitated because because Max is a strutted airplane, and only in the inverted position, the struts will carry loads ( in tension ?) like in normal flight ?
I am interested in the subject of load testing. It really is unbelievable the amount of load a wing can carry, when you see so many sand bags piled up atop the wing.
A test with the load applied on top of the wing has to be a negative g test. Usually, the negative g loads are much less except for aerobatic planes. For the Minimax, the wing structure can certainly handle the required negative g loads. The only question would be can the struts take the compression loads. I haven't run the numbers for the negative g loads. The stress report says the load for the main strut is 821 lbs. The equation for column buckling stress is P = π2 E I/ L2 where P is the critical buckling load, E is the elasticity modulus, I is the area moment of inertia and L is the length of the column.
I, for a thin-walled tube is π(r14 - r24)/4 where r1 and r2 are outer and inner radii.
E is 10,000,000 psi for aluminum. The strut length is 55", then the critical load is 1300 lbs.
I didn't do the covering, it was done before my time. The structure of the aircraft as a whole seems okay. The part of the wing right above the aileron cavity where the fabric glues to is separating from the ribs, but I don't think I can blame that on the builder. I've read multiple instances of older style minimaxes having that problem, and this one was ordered in 1990, so.. checks out.
Everything else seems glued on nicely. Fabric passed the puncture test with that special tool, so I'd hate to make more work for myself then necessary. I am thinking I'll stick a borescope down the wing to see the various joints, but other then that, I guess I have to trust it?
A test with the load applied on top of the wing has to be a negative g test. Usually, the negative g loads are much less except for aerobatic planes. For the Minimax, the wing structure can certainly handle the required negative g loads.
Quite possible the picture I saw was a negative 'g' load-test. Thanks.
When it is stated that wing loading is roughly 5 lbs per sq foot for the Max, does it mean that the wing fabric also carries some of the load ?
The load is the sum of the pressure distributed over the entire wing surface top and bottom. The fabric "feels" the pressure at every point and transfers the local loading to the wing structure. The load in the structure is a reaction to the fabric pressure. So one could say the fabric carries all the load into the ribs, spars, etc. Wing loading is more related to performance than structural concerns.
The load is the sum of the pressure distributed over the entire wing surface top and bottom. The fabric "feels" the pressure at every point and transfers the local loading to the wing structure. The load in the structure is a reaction to the fabric pressure. So one could say the fabric carries all the load into the ribs, spars, etc. Wing loading is more related to performance than structural concerns.
Isn't the wing covering a part of the overall structure design? Does it not contribute to the stability of the overall structure?
It does. However, I don't believe, for fabric covered wings, the contribution the covering makes is taken into account. For instance, the fabric certainly would help resisting wing bending but the spars are designed to take all the bending. I wouldn't want to rely on a perfect covering job (and for it to remain perfect!) for the integrity of my wings. It's enough for the fabric to take the air pressure. Stressed skin design is a different category altogether.
If testing with sandbags, is the covering not necessary to correctly allow the lumpen sandbags to simulate a real load?
Plywood sheets would help simulate a distributed load while providing some load themselves. The lift truss is composed of the spars and struts and the fuselage components that react those loads. So it would be entirely adequate to hang weights from points on the spars rather than pile weight on top. A rib can be tested separately, to destruction if desired, without a large investment.
Apologies for reopening an old thread.... I noticed the stress report does not cover (at least explicitly) the combined compression and beam loading on the inboard (of the strut) section of the wing. ANC 18 gives a method to do this. Any thoughts in why this would be?
I found the answer. Bob Daly clarified on another forum his interpretation of how compression of the inboard spar is in fact taken into account in the allowable BM in figure C2. In case anyone is interested ....essentially the max compression is transformed into an equivalent bending moment by fc*sectionModulus and subtracted from the total allowable bending moment.
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