Aerodynamic Design of the Bleed Slot in a Hypersonic Quiet Nozzle

The bleed slot is necessary for the requirement of the hypersonic quiet flow all over the world. The aim of the bleed slot is to decrease the influence of the disturbances from the contraction of the quiet nozzle to the boundary layer downstream of the throat, so that the boundary layer of the nozzle could be maintained as laminar flow. The main parameters of the bleed slot include the distance from lip to throat (DLT) and the width of slot (WS). Various values of those parameters will affect the performance of the slot by changing the suction intensity of the bleed slot. Two kinds of the bleed slots in the world are compared in this paper and the aerodynamic design of the bleed slots is optimized based on the Purdue-type slot. The influences of the various values of those parameters to the flow field around the throat are analyzed and the optimizing results of DLT and WS are consistent with those relative data designed for the slot of the Boeing/AFOSR Ma 6 Quiet Tunnel.

The thermochemical non-equilibrium scale effects of the high enthalpy nozzle

The high enthalpy nozzle converts the high enthalpy stagnation gas into the hypervelocity free flow.The flow region of the high enthalpy nozzle consists of three parts:an equilibrium region upstream of the throat,a non-equilibrium region near the throat,and a frozen region downstream of the throat.Here we propose to consider the thermochemical non-equilibrium scale effects in the high enthalpy nozzle.By numerically solving axisymmetric compressible Navier-Stokes equations coupling with Park’s two-temperature model,the fully non-equilibrium solution is employed throughout the entire nozzle.Calculations are performed at different stagnation conditions with the different absolute scales and expansion ratio.The results of this study are twofold.Firstly,as the absolute scale and expansion ratio increase,the freezing position is delayed,and the flow approaches equilibrium.Secondly,the vibrational temperature and Mach number decrease with the increase in the nozzle scale and expansion ratio,while the speed of sound,static pressure,and translational temperature increase as the nozzle scale and expansion ratio increase.

Integrated supersonic wind tunnel nozzle

In supersonic wind tunnels, the airflow at the exit of a convergent-divergent nozzle is affected by the connection between the nozzle and test section, because the connection is a source of disturbance for supersonic flow and the source of disturbance generated by this disturbance propagates downstream. In order to avoid the disturbance, the test can only be carried out in the rhombus area. However, for the supersonic nozzle, the rhombus region is small, limiting the size and attitude angle of the test model. An integrated supersonic nozzle is a nozzle and a test section as a whole, which is designed to weaken or eliminate the disturbance. The inviscid contour of the supersonic nozzle is based on the method of characteristics. A new curve is formed by the smooth connection between the inviscid contour and test section, and the boundary layer is corrected for the overall curve. Integrated supersonic nozzles with Mach number 1.5 and 2 are designed, which are based on this method. The flow field is validated by numerical and experimental results. The results of the study highlight the importance of the connection about the nozzle outlet and test section. They clearly show that the wave system does not exist at the exit of the supersonic nozzle, and the flow field is uniform throughout the test section.