抄録
This paper presents numerical investigations for unsteady flow fields in supersonic turbine stages for a Rocket Engine. Two dimensional analyses were conducted using spatial high-order scheme CFD code "Numerical Turbine". Analysis cases were single stage cases with two back pressure conditions, 1.5 stages case and 2 stages case. Results indicate pressure increase toward LE at suction side of 1^<st> Rotor blade. This pressure increase is caused because rotor blade passes through nozzle TE direct shock wave and reflect shock wave. Furthermore, vortex is created at near rotor blade suction side and is convected near suction side. As a result, pressure decrease toward downstream at suction side of 1^<st> Rotor blade occurs. Influences of back pressure are limited at rotor outlet region due to flow chokes at rotor passage near outlet.
