AERO2363 - Design of Aerospace Composite Structures - Engineering Assignment Help

Assignment Task

• Due date: Sun 20 September 2020, 11:45 PM, submitted online at AERO2363 Canvas via myRMIT.

• Individual assignment: One per student (multiple attempts allowed, but only last submission graded).

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Note:

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Use lamina properties given in Table 1 for all problems in this assignment.

Problem 1

You have been assigned with the design of a composite shell for the wing of PC9 aircraft. Under Design Limit Load (DLL), the composite laminate is subjected to given Nx, Ny and Nxy. The laminate configuration based on the aerospace design guidelines is required. Assume that zero and 90-degree plies do not contribute towards the shear strength of the laminate. 

Part A: Determining failure properties 

In-plane material failure properties for a single ply of IM7/977-3 unidirectional pre-preg need to be calculated. Using the 10% rule, determine the following: A1. Longitudinal tensile strength along A2. Longitudinal compressive strength A3. Transverse tensile strength along A4. Transverse compression strength A5. In-plane shear strength (S±45) 

0°, ±45° and along 0°, ±450°, ±45° and 90along 0°, ±45° 90° and ° (X90t0,X° (Xt±45and °(Yt090,Y°t±45(Yc0c0,X,Xc±45t90) 

,Xc90,Y,Ytc90±45) ,Yc90) 

) 

Part B: Preliminary laminate configuration 

Using the failure properties from Part A, estimate the minimum number of plies required in the main directions: N0,N45 and N90. 

RMIT Classification: Trusted 

Part C: Final laminate configuration 

For any aerospace composite structure, we need to follow the following design guidelines: 

• Use a minimum of 10% of plies in each of the principal directions [0,90, ±45] 

• Use balanced/symmetrical laminates to avoid warping 

• Use a maximum of four adjacent plies in any one direction 

• Place ±45°plies on the outside surfaces of shear panels 

• For thicker laminates (>16 plies) adjacent UD plies should be oriented with no more than 60° between them C1. Apply these design guidelines and finalise the number of plies in the main direction: 

N0,N45 and N90. C2. Report the laminate configuration. 

Part D: Control 

After finalising the laminate configuration, now it is the time to check the load-bearing capacity of the laminate. D1. Use eLaminate spreadsheet and calculate the in-plane properties of the finalised laminate: Ex, 

Ey,Gxy,νxy and νyx. D2. Use the unnotched failure strain for tension and compression (given in Table 1) and applied load (Nx, Ny and Nxy) in eLaminate spreadsheet to calculate the minimum Margin of Safety (M.S.) from Laminate-Based criterion. D3. Comment on the safety of the designed laminate. 

Part E: Notched strength 

Circular holes may be introduced in the designed laminate for bonding to stringers. The stress concentration factor and notched strength of the laminate should be analysed. E1. Using the in-plane properties from D1, calculate the stress concentration factor, kσ. E2. The Point Stress Criterion (PSC) is a simple empirical formulation to calculate the notched strength of composite laminates. Assume that the characteristic length for this laminate is do = 0.5 mm. Determine the largest circular hole radius where the notched strength of the laminate is 40% of its un-notched strength. 

Part F: What-if analysis 

During the service time, the composite laminate can be subjected to Design Ultimate Load (DUL), where DUL = DLL × 1.5. As a final check, calculate the margin of safety for Nx, Ny and Nxy considering the laminate configuration from C2 when it is subjected to DUL. 

RMIT Classification: Trusted 

Conceptual questions: 

State whether each of the following statements is True (T) or False (F): 

Q1. Since carbon fibre and epoxy are brittle materials with low fracture toughness values of 

1.6 MPa√m and 0.5 MPa√m, respectively, therefore, the carbon/epoxy composite is also brittle. Q2. In sandwich laminates under bending, the core supports shear stress while skin is under 

normal strain. Q3. Material properties in a composite laminate is a function of fibre orientation only. Q4. If the fibre orientation in a ply change from 45o to 90o, its tensile stiffness changes 

significantly. Q5. A composite ply exhibits maximum shear modulus when fibres are oriented along 45o 

concerning the loading direction. Q6. A [0/45/90/90/45/0], made of IM7/977-3 carbon/epoxy composite can be categorised as a 

quasi-isotropic laminate. Q7. The stress distribution through the thickness of a cross-ply laminate is not continuous. Q8. A16 in an extensional stiffness matrix member is responsible for shear-extensional coupling. Q9. D11 in a bending stiffness matrix is responsible for bending-torsion coupling. Q10. In a symmetric laminate, D16 from the bending stiffness matrix is always zero. Q11. If the longitudinal fibre tensile and compressive strength of IM7/977-3 unidirectional pre- 

preg is 2,825 MPa and 2,275 MPa, respectively, its in-plane shear strength is about 114 MPa. Q12. The extensional modulus of a [±45]4S laminate, made of IM7/977-3 unidirectional pre-preg 

is about 16.4 GPa. Q13. The allowable tensile strength of a [903/±452/03]S laminate is higher than that of a 

[904/±452/03]S laminate. Q14. The modulus of a [03/±452/902]S laminate is higher than that of a [04/±452/903]S laminate. Q15. The Design Ultimate Strain of a composite laminate is 30% of its allowable strain. Q16. The Design Limit Strain of a composite laminate is 30% of its Design Ultimate Strain. Q17. Aerospace composite laminates should be symmetric to avoid extension and shear coupling. Q18. Aerospace composite laminates should be balanced to avoid bending and twisting coupling. Q19. A component made of IM7/977-3 carbon/epoxy laminate is subjected to Nx and Ny only. 

Since there is no load along 45°a cross-ply laminate configuration can be used for this component. Q20. Following the aerospace composite design guidelines, an [45/0/90/-45/0/90/45/-45/90/0]S 

laminate is an acceptable stacking sequence.

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