Investigation Of Supersonic Non-uniform Flow Simulation Method

Integration design is required by hypersonic vehicle and propulsion system, as a result the lower energy flow of vehicle forebody boundary layer will enter the inlet. Because of the non-uniform of the inlet exit flow and the peculiarities of the propulsion system inner flow, the incoming flow of every propulsion system assembly is obviously non-uniform. In order to study every propulsion system assembly in the free jet experiment and the direct-connect experiment, supersonic non-uniform flow fields with special distributions are required as reference incoming flow fields.The primary content of the thesis is to form supersonic non-uniform flow fields with special distributions, through viscous friction, curved shock, oblique shock and expansion wave, based on basic properties of hypersonic and supersonic flow.Firstly, investigations of Mach number, static pressure and flow direction angle at different development section positions after the exit of the two dimension plane array viscous passage, three dimension square louver array viscous passage and three dimension regular hexagon honeycomb louver array viscous passage are carried out, which indicate that flow parameters distributions are zigzag locally, and have overall space gradient. The local serration degree and space gradient can be weaken as the development section length increased, due to the enhanced mixing. However, as a simple way to generate non-uniform flow, it is still worth to be considered.Secondly, the wall surface shape can be designed, the curved shock shape can be formed and the flow with expected non-uniform Mach number distribution could be obtained, availing the rotational flow characteristic line method, according the flow parameters behind the curved shock. And the conclusion, that the curved shock shape is controlled by a very short wall surface, is got at the same time. The two dimension blunt body is designed by this method, and the non-uniform supersonic flow, whose Mach number distribution is power law, could be produced by curved shock. The wind tunnel test indicates that, in principle, it is a feasible approach to produce a non-uniform supersonic flow.Moreover, a method to create supersonic flow that can change Mach number continuously was investigated. A simple device, that can form continuously changed Mach number flow only by adjusting the wedge angle, was designed availing the theory of shock wave reflection. With the Mach 3.85 coming flow as an example, the method was theoretically proved to be possible.In the end, availing the rotational flow characteristic line method, consulting wind tunnel nozzle design method, the core section curve and the transition section curve of the nozzle are designed. The precision of rotational flow characteristic line method is validated availing the uniform incoming flow and the linear distribution non-uniform incoming flow, which indicates that the relative error of the characteristic line method declines as the increase of point numbers in initiation area. Through an isentropic expansion surface specially designed for the coming flow of supersonic boundary layer flow developed from a flat plane, low Mach number supersonic non-uniform flow can be accelerated isentropically into high Mach number, and at the same time the range of non-uniform region can be enlarged, i.e. a large region of non-uniform supersonic flow can be obtained in a short distance.