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Preliminary design of a composite fuselage?

Preliminary design of a composite fuselage?

Preliminary design of a composite fuselage?

I've just started studying composites design. To get a better idea of how the stuff I learn fits into the big picture, it'd be great if someone could give me just a brief overview of the process of designing a fuselage with an arbitrary shape (not necessarily a simple circle).

1) Starting from scratch for the design work but supposing I know the sorts of loads applied (lifting loads on the wing for example) and the geometry of the fuselage, how would I obtain a preliminary composite layup so I can start somewhere? Do I create a preliminary composite layup in a FEA software and analyze it?
2) If analyzing by hand (might be necessary for preliminary design in which the composite layup is not yet known), is analyzing the structure done with some sort of method to transform the loads and stresses as might be necessary due to the various curvature that might occur in a fuselage such as this sort? What's the proper term for this?

RE: Preliminary design of a composite fuselage?

The question is such that a fairly long reply is the only option I can see, so apologies.

A fuse being a shell you have a basic decision to make between space frame with sheet or fabric skin, monocoque, and semi-monocoque. Many smaller aircraft use a sandwich shell structure for the fuselage.

First take predictions of torsion, shear and bending along the fuselage. These may come from aerodynamic sources or heavy landings or dynamic cases such a losing part of a propeller blade. Five diagrams of maximum load should be predicted for each of torsion, X and Y bending and X and Y shear vs. longitudinal position.

For a sandwich structure, failure can be due to shear (mainly from torsion cases, but the maximum from the combination of torsion and X and Y shear must be used), compression and tension from bending.

To check failures you will need compressive skin modulus (Es) and skin shear modulus (Gs) and core through-thickness modulus and also flatwise strength in tension and compression (Etcore, Eccore, Stcore and Sccore). The through-thickness properties are largely to check for skin wrinkling. You will also need to check overall compression buckling and shear buckling (needs compressive and shear moduli again) and overall tension strength, compression strength and shear strength as well (needs three skin material strengths, St, Sc, S; core, even if it's foam, can be conservatively ignored).

There's not much inside the fuse to take up space (mainly control cable runs for rudder and elevator; not sure where you intend to put the fuel tank) so the core thickness can be quite large. Pick a sensible thickness (probably and inch or two) and don't worry about adjusting it with the length dimension.

You also need to pick an initial skin material. Although most aircraft of the type your illustration shows have used woven glass fiber skins (composite density 2000 kg/m^3), a woven carbon fiber is another option (composite density 1560 kg/m^3). With carbon skins and allowables for damage you can ignore fatigue. If you use glass you should probably do some fatigue analysis (more loads needed).

As initial values for carbon fiber skins take Es 50 GPa (7.25 Msi), Gs 15 GPa (2.2 Msi), St 345 MPa (50.0 ksi), Sc 175 MPa (25.4 ksi), S 110 MPa (16.0 ksi). Note: these are values for woven "HS" material with a layup of "50:50" and allow for impact damage causing a dent up to 1.25 mm deep. (HS is the type of fiber and 50:50 means an equal number of plies at 0° and 45°.) It is possible to use "IM" fiber rather than HS but this is about 20% better and is two or three times more expensive.

Or, as initial values for woven glass fiber skins use Ec 22 GPa (3.2 Msi), Gc 8.5 GPa (1.2 Msi), St 200 MPa (29 ksi), Sc 170 MPa (25.0 ksi), S 100 MPa (14.5 ksi).

All these numbers I've supplied will be fairly controversial. You should of course have validated allowables for the chosen materials with appropriate damage rather than these guesses.

Using a 3 lb core (3 lb/cu.ft, or 48 kg/m^3) should give you adequate manufacturability and properties. Use core strength and moduli from Hexcel (honeycomb, use 1/8" cell size Nomex material) or Gurit or Rohacell for foam.

These moduli can be used for predicting onset of buckling stresses according to Roark, Bruhn, NASA CR 912 and NASA CR 1457. Buckling equations for composite plates and cylinders can be used and these are slightly more complicated than the ones in Roark and Bruhn but are similar in form.

Modify the material properties to account for actual skin thickness required, using classical laminate theory to predict properties for layups which are not 50:50. If using FE analysis then this can be largely automated. It is possible that properties of skins will be arrived at by test which would eliminate the calculation part of this. If composite buckling formulas are being used then these will have to be revised in line with the layup used in each part of the fuse.

Do trade-off studies to see what effect core thickness has on weight and minimize core thickness accordingly. You should find that skin wrinkling is critical for areas of high compression and shear and this is affected by all of core strength, stiffness and thickness, but the basic skin strength may still be critical. This skin strength should not be much affected by core considerations. If the core trade-offs favor a thin core the overall stability may become critical. Using a thicker core should avoid this for a minimal weight penalty.

That's about the simplest description I can come up with, twenty or thirty lines. The main design complication compared with metal skins is the adjustment of properties for different skin thicknesses. The main design simplification of carbon composite compared with metal is that its fatigue resistance largely obviates any fatigue analysis.

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