Air inlet channel shock wave boundary layer interference characteristic measurement method

Abstract

The invention discloses an air inlet channel shock wave boundary layer interference characteristic measurement method. Firstly, a scaling frame type test model is designed for an engine air inlet channel according to a wind tunnel test simulation similarity criterion; pressure-sensitive paint is sprayed in a to-be-measured area of the scaling frame type test model, and a PSP measurement system is constructed for measuring the pressure distribution characteristic of the inner surface of the air inlet channel, the size and position of shock waves, the flow state development of a shock wave boundary layer and the distribution characteristic of a surface vortex approximate friction force line; a PIV measurement system is established for measuring velocity field information and a wave system structure of a flow field in the air inlet; and a schlieren measurement system is built and used for measuring laminar flow turbulence flow state changes and boundary layer thickness in the air inlet channel. According to the method, the defects of parameter measurement of an existing air inlet channel experiment project can be effectively overcome, and a global and visual identification method is provided for typical shock wave boundary layer interference flow structures, separation characteristics and shock wave series flow fields.

Description

technical field [0001] The invention belongs to the technical field of aerospace experiments, and relates to a method for measuring the interference characteristics of an inlet shock wave boundary layer. Background technique [0002] Shock boundary layer interference is a complex flow phenomenon commonly found in super / hypersonic aircraft, which often has a significant impact on the aerodynamic / thermal distribution of the aircraft, thereby affecting the aerodynamic performance and heat protection characteristics of the entire aircraft . Especially for the analysis of shock boundary layer interference characteristics of internal flow fields such as transonic wings and cascades, scramjet inlets and isolation sections under complex incoming flow conditions, there is a lack of wave system structure and flow analysis for complex flow fields. The comprehensive observation and analysis of unsteady characteristics of multi-physics parameters such as state changes and full-field dyn...

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Problems solved by technology

Especially for the analysis of shock boundary layer interference characteristics of internal flow fields such as transonic wings and cascades, scramjet inlets and isolation sections under complex incoming flow conditions, there is a lack of wave system structure and flow analysis for complex flow fields. The comprehensive observation and analysis of unsteady characteristics of multi-physics parameters such as state changes and full-field dynamic pressure changes, while the existing research methods lack the overall and comprehensive analysis of the typical shock boundary layer disturbance flow structure, separation characteristics, and shock string flow field. intuitive identification method

The traditional measurement method generally uses a single-point sensor to measure the internal pressure characteristics of the engine to predict the flow separation. Since there cannot be too many points, the accuracy of judging the separation position is very low, and it is powerless for large-area pressure fluctuations. The result is very unintuitive and the amount of information obtained few

Most of the inlet experimental projects are limited to the measurement of the three parameters of the total pressure recovery coefficient, the flow coefficient, and the uniformity of the flow field at the outlet of the inlet. , laminar flow, turbulent flow state change, boundary layer thickness, shock boundary layer three-dimensional wall effect, surface eddy current approximate friction force line distribution and other internal flow characteristics cannot be effectively observed

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