ガスタービン翼の熱疲労

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Leading edge cracks on blading of aircraft jet-engines have been found out frequently and systematic investigations are continued in US Air Force and N.G.T.E. in England. Thermal fatigue due to repeated thermal stress induced in transient engine condition is the dominant cause of such crackings. To investigate thermal fatigue behaviour, we estimated local heat transfer around blade profile and calculated thermal stress distributions in transient condition. Experiments were carried out under the similar condition to engine start and stop, using model blades, and investigated origination and propagation os such crackings. It is essentially important to investigate heat transfer around blade profile to estimate temperature distribution in blade section. We applied the approximate method shown by E.R.G. Eckert in N.A.D.A. technical report in 1955. Transient temperature distributions were estimated by using Capacitance-Resistance net work simulator and stress distributions were calculated assuming young's moculus and linear expansion coefficient as temperature dependant by using an electronic digital computer. We concluded that cyclic plastic strain must have been occurred in start and stop of gas turbine. 4 different materials N-155, S-816, 25Cr-20 Ni and 18Cr-8 Ni were investigated in the thermal fatigue test. Model blades were heated rapidly by hot gas stream and then cooled by cold air, and origination and propagation of crackings were investigated. The conclusions of our study for improvement of blade begavior toward thermal fatigue are as follows : (1) Leading edge radius should be as large as possible within the allowable limit of aero-dynamics. (2) Air film cooling injecting air from leading edge will be effective. (3) Experiments showed slits to absorb thermal strain at leading edge are effective. (4) Material of higher thermal fatigue resistibility is preferable.
1963-03-15

ガスタービン翼の熱疲労

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