By Baker A.A.
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Additional resources for Advances In The Bonded Composite Repair Of Metallic Aircraft Structure
At this point no more crack growth occurs due to the proximity of the panel boundary. The only growth that occurs is at the right hand crack tip. In this study the maximum crack length considered is 196mm. 1. -repaired mode 2 -A-repaired mode 3 _ _ A 0 & > -- ‘I--=-=:- $ $ 400- A -0 300: 200100- o - n . , . , . , . , . , . , . , . , - Fig. 9. Frequency versus crack length for un-repaired and repaired cases Chapter 19. 1 Hz 54 1 Fig. 10. Mode shapes for the repaired panel containing a short crack (2a=50mm).
35 as a function of temperature. The service temperature for this adhesive ranges from -55 "C to 150 "C. Only two data points exist for each temperature case, as a result values for shear modulus have been extrapolated. 419T (MPa), where Tis the temperature in "C. - 20°C c 40°C > a, IO2 d- E LL --P- 10' 60"C 80°C 100°C + 120°C + 140°C 1o o IO-' loo 10' lo2 lo3 l o 4 l o 5 lo6 lo7 IO* lo9 Scale Fig. 34. Dyad 606 material data. (To obtain loss factor divide shear modulus scale by IO6). Advances in the bonded composite repair of metallic aircraft structure 562 Loss factor Shear modulus (D .
2dB), [l]. Chapter 19. 1 disturbance is responsible for the higher than expected SPL. The possibility exists that the flap may be responsible for this. In the case of the B52, , it was found that cracking in the aft fuselage skin was a result of the aerodynamic disturbance caused by the deflected flap position during take-off. However aerodynamic disturbances can also be caused by high angle of attack manoeuvres resulting in separation of airflow over the flap. 3. s. e. 1. 4. Random response analysis The random response analysis capability of the NASTRAN program has been used to solve this problem [Ill.