Shockwave-induced boundary layer bleed for transonic gas...

Details Shockwave-induced boundary layer bleed for transonic gas... Shockwave-induced boundary layer bleed for transonic gas...

F - Mech Eng,Light,Heat,Weapons

02

C

F02C 9/18 (2006.01) F01D 5/18 (2006.01) F02C 3/06 (2006.01)

Patent

CA 2565867

An apparatus and method is provided for improving efficiency of a transonic gas turbine engine compressor (20) by bleeding off a shockwave-induced boundary layer from the gas flow passage (42) of the compressor using an array of bleed holes (36) having a downstream edge (39) aligned with a foot (45) of an oblique shock wave (44) which originates on the leading edge (46) of an adjacent transonic rotor blade tip (30).

L'invention concerne un appareil et un procédé permettant d'améliorer l'efficacité d'un compresseur (20) de turbine à gaz transsonique par prélèvement d'une couche limite induite par une onde de choc depuis un passage d'écoulement de gaz (42) du compresseur au moyen d'un réseau d'orifices de prélèvement (36) présentant un bord aval (39) aligné sur un pied (45) d'une onde de choc oblique (44) formée sur le bord d'attaque (46) d'un embout (30) d'aube de rotor transsonique adjacent.

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