Aircraft Loads Simplification

[alex2000junio]

I'm working on my final college project on composite materials and studying one of the standard models provided by NASA CRM. I'll be conducting a buckling study in FEMAP on the upper wing surface.
To do this, I first need to apply the loads. For the loads, I'm doing the following: I'll consider and calculate the trim state (calculating the lift, thrust, drag, and weight so that their sum equals zero). For the force distribution, I'll use the Stender (Schrenk) method. I only need to consider the wing loads, as the rest of the aircraft model won't be developed.
Since the focus of my work is composite materials and their behavior under buckling, I'm considering this simplification and using these loads and this distribution method. Although I have experience with simulation, I don't have it with aircraft loads, and I wanted to know if considering only these loads is a valid simplification. If not, could you please point me to the load I have to consider and maybe some bibliographical references, as I'm a bit lost.

Any help I'll appreciate!

 

[rb1957]

if your project is about the buckling of the skin, and the modelling the FeMap, then that's what to focus on ... not the flight loading to create a stress on the wing panel. Just assume a stress level.

 

[sdm919]

I agree with rb1957's comments. Skin buckling and wing loads are more like two separate and large subject areas. You may be making your study too complicated by combining them.

If your study is more about buckling of the skin panels, you could assume ranges of loads as rb1957 suggested. Axial compression loads Nx values in the range of perhaps 1000-10000 lb/in. One set with zero shear (Nxy=0) and another with a value of shear a certain fraction of Nx, say Nxy = 20% Nx. One of the capabilities of composites is getting different behavior by varying the ply angles. So your study could be about that.

If your study is more about modeling wing loads, you could make some simple calculations based on classic methods. Basically treating the wing as a cantilever beam and computing the vertical load distribution, then from that the wing shear, bending and torsional loads at stations along the span. One reference is: "Aircraft Structures" by David J. Peery (the original 1950 edition, or the 2011 reprint by Dover Publications, not the 1982 edition by Peery & Azar).

Chapter 9 of Peery discusses "Spanwise Air-Load Distribution" and Chapter 10 discusses "External Loads on the Airplane".