Design And Mechanism Investigation Of Contour Of Supersonic Inner Flow-path

The studies of supersonic inner flow-path are crucial in the design of high-speed aerospace vehicles and weapons system,which are also important for understanding the mechanism of the boundary layer transition,shock characteristics,shock wave/shock wave interaction,shock wave/boundary layer interaction,boundary layer separation induced by the inverse pressure,and shock-train behavior.In the inner flow-path,the flow has high speed and three-dimensional and highly unsteady natures,so the flowfield with free-shock wave is difficult to be achieved,which presents special challenges for the studies of the supersonic inner flow-path and has been becoming a research hotspot but difficulty in the study of the supersonic aerodynamic.The thesis reviewed the research status of the design method of the supersonic nozzle and the hypersonic inlet,and analyzed the research situation of the shock train in the restricted duct.It is concluded that the present aerodynamic design method of the supersonic contour primarily focused on the passive adjustment.This method requires the given flowfiled boundary in advance to achieve the flow parameters,which is rather complicated and time consuming.The research of the shock train in the restricted duct focused on the straight flow-path or slightly-expansion duct.The research of the oscillation flow in the restricted duct focused on the over-expended nozzle,which aimed at the single shock wave,rather than the shock train.Based on the research of the supersonic inner flow-path in China and abroad,the thesis conducted the investigation of the inverse aerodynamic design method and the mechanism of the shock train in the S-shaped flow-path.In order to achieve multi-Mach number running for the present wind tunnel,a straightforward technique has been developed to quickly determine a design of the super/hypersonic nozzles sharing throat section with multi-Mach number exiting flow.The simulation results indicate that the good flow quality is attained for three nozzles within ?Ma/ Ma < 1% in the test rhombus region.A straightforward technique has been developed to quickly determine a design of the super/hypersonic dual-inflection nozzle.The simulation results indicate that the good flow quality is attained for different nozzles within Mach number error less than 1.2% in the test rhombus region.A design method has been developed to specify the contour of the continuously variable Mach-number nozzle for the supersonic wind tunnel.The computed results indicate that exit uniform flow was obtained with 1.19% of maximal Mach-number deviation at nozzle exit.A design method of planar curved shock waves has been developed.The curve of shock wave is specified in advance with the B-spline curve and the domain of influence of the shock wave and the wall determining the shock were solved by the rotational method of characteristics.With inviscid CFD simulation of the concave shock,the present design method has been validated.Three types of shock wave,including straight shock,concave shock and convex shock,were designed and compression capacity,compression efficiency,capacity of capturing mass flow,and the capacity of turning the flow were investigated.On design point,the principle of contraction ratio,total pressure recovery coefficient,pressure rise,angle of flow turning versus curvature of the shock wave has been given.Meanwhile,on off-design point,the principle of the total pressure recovery coefficient and flow mass coefficient versus angles of attack and Mach numbers has been shown.With centerline Mach-number distributions provided in advanced,two-dimensional S-shaped supersonic shock-free duct has been designed by the method of characteristics.The relationship between the shock train and back pressure was investigated,which indicated that it was difficult for the head of the shock train to be steady in the vicinity of the duct inflection.The flowfield had dual solutions when the head of the shock train was in the vicinity of the duct inflection,that is,for the same duct and the given same back pressure and different initial flowfield,two different convergent flowfileds were obtained.It is indicated that the flow hysteresis exists in the S-shaped duct.Meanwhile,the borderlines of hysteresis quadrangle have been analyzed.In the hysteresis loop,when the border point evolves into the bifurcate point,that is,dual-solution domain evolves into single-solution domain,the larger separated region in the vicinity of wall transfers quickly from one wall to the other and shock-train structure evolves.The structure and dynamic characteristics of the shock train in a supersonic S-type duct have been investigated experimentally.The experimental results indicate that the higher the back pressure,the lower the oscillation frequency of the shock train,but axial oscillation distance of the shock train increases at first,and then decrease.When the blocking ratio is 53.54%,the large separation zone induced by the inverse pressure appears in the vicinity of the top wall,and the special shock train forms in the S-shaped duct.The dominant frequency of the shock train is 426 Hz,and maximal axial oscillation distance is 0.6 times larger than the duct height.When the blocking ratio is 64.65%,the large separation zone appears in the vicinity of the bottom wall,and the Mach reflection and regular reflection of the separation shock wave alternately appear.The dominant frequency of the shock train is 246 Hz,and the maximal axial oscillation distance is 2.7 times larger than the duct height.When the blocking ratio is 68.60%,the trend of the location of the shock train with the time has no special characteristics and the periodicity is not obvious.The maximal axial oscillation distance is 0.9 times larger than the duct height.