Investigation On The Organization Method Of Passage Shock Wave And Control Of Leading-Edge Passage Shock Wave/Boundary-Layer Interaction Within “Shock-in-Type” Supersonic Compressor

With the continuous improvement of aero-engine performance requirements,the aero-engine has been promoted towards the direction of high thrust-weight ratio,low fuel consumption,high maneuverability and high reliability.As an important index to evaluate the performance of aero-engine,thrust-weight ratio has considerable influence on the flight speed and maneuverability of aircraft.Compressor is one of the key parts of aero-engine,its length and weight account for about half of the whole engine.Therefore,it is crucial to improve the thrust-weight ratio of aero-engine by improving the load of the compressor stage,reducing the number of compressor stages and developing a compact aerodynamic layout.With increasing the load of compressor stage,the Mach number at the rotor inlet has been also increasing,and the relative supersonic flow appears.Under the relative supersonic operating conditions,the organization of shock system in the compressor rotor blade passage and the corresponding losses brought about by the strong shock and shock-induced boundary layer separation,significantly influence the overall compressor aerodynamic performance.Therefore,it has become a key scientific problem in the design of supersonic compressor to reasonable organize the shock wave system structure in the blade passage and reduce the flow separation loss induced by the interaction between the strong shock wave root and the boundary layer while using the shock wave pressurizing in the main zone of the blade passage,exploring the potential of efficient pressurization by using the shock wave.Based on the supersonic compressor cascade,analyses on the internal relationship between the shock wave system and pressurizing characteristics within supersonic blade passage have been implemented,exploring the physical mechanism of the effect brought about by local blade curvature variation on the boundary layer separation induced by leading-edge passage shock,a design method for negative curvature profile with constant adverse pressure gradient was proposed to suppress shock-induced boundary separation in supersonic compressor cascade.Furtherly combined the three-dimensional flow characteristics in supersonic compressor rotor,the boundary layer separation suppression method by local blade profile negative curvature was developed of supersonic compressor rotor.(1)Based on the supersonic compressor with design inlet Mach number of 1.75,the organization method of passage shock wave and analysis of the internal relationship between the shock wave system and pressurizing characteristics within supersonic compressor cascade have been implemented,illustrating the local boundary layer separation structure characteristics induced by strong adverse pressure gradient originating the shock foot of the leading-edge passage shock,and distribution characteristics of blade load in the interaction region between shock wave and boundary layer as well as the influencing factors of shock wave-induced boundary layer separation are clarified,which laid a foundation for the subsequent research of flow separation suppression methods.(2)Investigation on supersonic compressor cascade experimental test based on supersonic compressor cascade test bench has been carried out.The pressurization potential of shock wave system was verified by pressure test and the structure of leading-edge passage shock foot was further analyzed by schlieren photographs,which clarifies the relationship between the load distribution characteristics on the blade surface and shock wave structure within blade passage.In addition,the accuracy of the numerical calculation method was examined,laying a foundation for the subsequent research work.(3)Based on the boundary layer separation structure captured by numerical method and experimental test induced by the leading-edge passage shock wave within supersonic compressor cascade,a design method on the basis of local negative curvature with constant adverse pressure gradient for blade profile to suppress shock-induced flow separation has been proposed.The influence mechanism of local negative curvature blade profile modification upstream of the leading-edge passage shock root at the blade suction surface on the shock wave root structure,the blade load distribution characteristics,the boundary layer behavior and entropy production variation across the interaction region between shock wave and boundary layer in blade passage are analyzed in detail.The design method of local constant adverse pressure gradient negative curvature blade profile and the shock-induced flow separation suppression method with blade curvature in supersonic compressor cascade are established.(4)Based on the blade curvature suppression method for shock-induced boundary layer separation in supersonic compressor cascade,the three-dimensional flow characteristics in the blade passage of supersonic compressor rotor were further take into account.It clarifies the action mechanism of local negative curvature profile upstream of the shock wave root of three-dimensional compressor rotor blade on shock wave-induced flow separation and radial transport of low momentum boundary layer fluid,and reveals the influence law of negative curvature profile design on shock wave/boundary layer interaction under different shock wave intensities and incident positions.A shock induced boundary layer separation suppression method for supersonic compressor rotor is developed.In this study,based on the mechanism of local blade curvature variation on shock-induced boundary layer separation,a design method of local constant adverse pressure gradient negative curvature blade profile was proposed to effectively suppress the boundary layer separation induced by the leading-edge passage shock in supersonic compressor cascade.The local negative blade-curvature suppression method for shock-induced boundary layer separation was further applied from two-dimensional supersonic compressor cascade to three-dimensional rotor,formed the design method of local constant adverse pressure gradient negative curvature blade profile for suppressing shock-induced boundary layer separation of supersonic compressor rotor.Utilizing the pressurizing effect provided by the shock wave in the mainstream region of blade passage and reducing the losses caused by the shock-induced flow separation at the shock foot,the overall aerodynamic performance of supersonic compressor rotor is promoted,which provides support for the optimization design of the supersonic compressor rotor.