Yoshiaki Kakuta

Effect of Isothermal Aging on Ultimate Strength of High-Temperature Composite Materials for SST Structures

The objective of this experimental study was to evaluate the effect of iso-thermal aging on the ultimate strength of three kinds of carbon/high-temperature composite materials, i.e., G40-800/5260 and MR50K/MR2000N bismaleimide composites and T800H/PI-SP amorphous thermoplastic polyimide composite. These materials are current candidate structural materials for a supersonic transport of the next-generation. The hole-notched and unnotched panels, before being machined to specimens, were isothermally aged at 120°C and 180°C for up to 15,000 hours. Static tests at room and elevated temperatures before and after thermal aging provided the open-hole tensile, open-hole compressive, and short beam shear strengths. Moreover, the effects of five oxidation resistant treatments on open-hole compressive strength at 180°C were investigated after isothermal aging of 5,000 hours at 180°C. The test results clarified the effects of isothermal aging on ultimate strengths and oxidation resistant treatments on open-hole compressive strength. Moreover, the possibility of developing an accelerated aging test method is discussed using a modified Larson-Miller type equation.

G40-800/5260 Carbon Fiber/Bismaleimide Composite Material: High Temperature Characteristics of Static and Fatigue Strengths

A lot of polymer composite materials are being used in the structures of civil transport aircraft currently under development, such as the Boeing 787, Airbus A350, and Bombardier C Series. Their percentages of structural weight are announced at 50%, 53% and 46%, respectively. However, mostly carbon fiber/epoxy (CF/Ep) systems are being introduced into their primary structures and they are only being introduced in environments where they are not expected to encounter high temperature. On the other hand, application of carbon fiber/bismaleimide (CF/BMI) composite materials is being expanded, especially for military aircraft structures such as the airframes of the F-22 Raptor and the F-35 Lightning II Joint Strike Fighter, the jet engine nacelle skins of the F-35 as well as the thrust reverser structure of the Gulfstream G450 business jet in civil aircraft (McConnell, 2009), and is expected for structures of the next-generation supersonic transport (SST). There are several reasons for CF/BMI system application for the structures described above. As for epoxy system composite materials, about 70°C is usually set as the design limit temperature for long term use (Brandecker & Hilgert, 1988, Fawcett et al., 1997). Meanwhile, carbon fiber/polyimide (CF/PI) system composite materials can be used for hotter structures, although they are very expensive and involve complicated processes. CF/BMI composite materials offer temperature performance and costs between those of epoxy and polyimide systems. Moreover, CF/BMI systems can be easy handled in an airframe parts manufacturing process in a way that is equivalent to that for epoxy systems. The design limit temperature of CF/BMI composite materials for aircraft structures is supposed to be around 120°C on the basis of actual application to the mechanically loaded structures described above. If the design limit temperature is set as 120°C, it is necessary to know the detailed characteristics of static and fatigue strengths at about 150°C from the view point of the safety margin; moreover, 150°C is considered to be close to the servicelimit temperature for CF/BMI composite materials. Therefore, in order to apply CF/BMI composite materials for aircraft structures that encounter medium high temperatures,