To reduce pressure loss by controlling a shock wave generated at a leading edge of a high turn/high transonic aerofoil used for a blade cascade for an axial compressor.
A flow rate distribution at a backside of the leading edge of the high turn/high transonic aerofoil used for the blade cascade for the axial compressor has supersonic portions k to l which are located at a rear side of a first local maximum value j of the flow rate are approximately constant in flow rate and within a 15% position of a chord length. In the supersonic portions k to l, a value dividing a Mach number difference ΔM between the back and forth ends by a chord length direction ΔX/C is less than 1, and the maximum Mach number of the transonic portions k to l is less than 1.4. A great first shock wave is positively generated at the position where the flow rate becomes the first local maximum value j to weaken the second shock wave generated at the transonic portions k to l whose rear flow rate is approximately constant so that pressure loss in a rear flow of the blade can be greatly reduced by suppressing a boundary-layer separation owing to the second shock wave.
ARIMA TOSHIYUKI
MURATA YASUSHI
Kazuaki Niki
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