The high-load compressor plays an important role in further improving the performance of aero-engine.However,the complex shock waves in the cascade channel also bring more aerodynamic losses.This paper proposes a supe...The high-load compressor plays an important role in further improving the performance of aero-engine.However,the complex shock waves in the cascade channel also bring more aerodynamic losses.This paper proposes a supersonic compressor cascade modeling method based on the theory of unique inlet flow angle,and the aerodynamic design and optimization of a cascade with inlet Mach number 1.85 are studied by combining the numerical optimization method and planar cascade experiment.The results show that pressure increase can be achieved by multiple shock waves which are obtained by the reflection of the leading edge detached shock wave in the initial supersonic cascade channel at the design point,which verifies the feasibility of the design method.After optimization,the aerodynamic performance of the cascade has been improved to different degrees at the design point and off-design point.When the static pressure ratio is 3.285,the total pressure recovery coefficient reaches 86.82%at the design point,which is on the advanced level of the same type of cascade.The experimental results of planar cascade schlieren and surface pressure measurement also verify the correctness of the simulation method,which provides useful references for the subsequent compressor design.展开更多
The aim of this paper is to predict the phenomenon of laminar separation, transition and reattachment in a low-pressure turbine （LPT）. Self-developed large eddy simulation program of compressible N-S equations was u...The aim of this paper is to predict the phenomenon of laminar separation, transition and reattachment in a low-pressure turbine （LPT）. Self-developed large eddy simulation program of compressible N-S equations was used to describe the flow structures of T 106A LPT blade profile at Reynolds number of 1.1×10^5 based on the exit isentropic velocity and chord length. The com- putational results show the distributions of time-averaged wall-static pressure coefficient and mean skin-friction coefficient on the blade surface. The locations of laminar separation and reattachment points occur around 87% and 98% axial chord, which agree well with experiment data. The two-dimensional shear layer is gradually unstable along the downstream half of the suc- tion side as a result of the spanwise fluctuation and the roll up of shear layer via Kelvin-Helmholtz （KH） instability. Three-dimensional motions appear near 84% axial chord which later triggers spanwise vortexes and streamwise vortexes, leading to transition to turbulence in the separation bubble. Through introducing the concept of dissipation function, the high loss mainly comes from the places where strong shear layer and intense fluctuation exist. Furthermore, the separation region is only an accumulation center of the low-energy fluid rather than an area of loss source.展开更多
基金funded by the National Science and Technology Major Project(J2019-II-0016-0037).
文摘The high-load compressor plays an important role in further improving the performance of aero-engine.However,the complex shock waves in the cascade channel also bring more aerodynamic losses.This paper proposes a supersonic compressor cascade modeling method based on the theory of unique inlet flow angle,and the aerodynamic design and optimization of a cascade with inlet Mach number 1.85 are studied by combining the numerical optimization method and planar cascade experiment.The results show that pressure increase can be achieved by multiple shock waves which are obtained by the reflection of the leading edge detached shock wave in the initial supersonic cascade channel at the design point,which verifies the feasibility of the design method.After optimization,the aerodynamic performance of the cascade has been improved to different degrees at the design point and off-design point.When the static pressure ratio is 3.285,the total pressure recovery coefficient reaches 86.82%at the design point,which is on the advanced level of the same type of cascade.The experimental results of planar cascade schlieren and surface pressure measurement also verify the correctness of the simulation method,which provides useful references for the subsequent compressor design.
基金supported by the Foundation for Innovative Research Groups of the National Natural Science Foundation of China(Grant No.51121004)the National Natural Science Foundation of China(Grant No.50976026)
文摘The aim of this paper is to predict the phenomenon of laminar separation, transition and reattachment in a low-pressure turbine （LPT）. Self-developed large eddy simulation program of compressible N-S equations was used to describe the flow structures of T 106A LPT blade profile at Reynolds number of 1.1×10^5 based on the exit isentropic velocity and chord length. The com- putational results show the distributions of time-averaged wall-static pressure coefficient and mean skin-friction coefficient on the blade surface. The locations of laminar separation and reattachment points occur around 87% and 98% axial chord, which agree well with experiment data. The two-dimensional shear layer is gradually unstable along the downstream half of the suc- tion side as a result of the spanwise fluctuation and the roll up of shear layer via Kelvin-Helmholtz （KH） instability. Three-dimensional motions appear near 84% axial chord which later triggers spanwise vortexes and streamwise vortexes, leading to transition to turbulence in the separation bubble. Through introducing the concept of dissipation function, the high loss mainly comes from the places where strong shear layer and intense fluctuation exist. Furthermore, the separation region is only an accumulation center of the low-energy fluid rather than an area of loss source.
