Research On The Thermal Protection By Opposing Jet And Transpiration For Vehicle Leading Edge And The Complex Flow Algorithm

The opposing-jet and inspiration leading edge offers long-range hypersonic vehiclean active thermal protection method with simple structure, high reliability, low cost andhigh efficiency. The leading edge is built with plantlets, and has both transpirationcooling and opposing jet for thermal protection. When it works, the transpirationcooling structure will cool down the long-range hypersonic vehicle, and the opposing jetwill also reduce the drag while reducing the heat flux. At the same time, it can make thehypersonic vehicle work for a long time and reusable.The work in this thesis is about the thermal protection mechanism of theopposing-jet and inspiration leading edge for hypersonic vehicle, which includes twoparts: one is opposing jet thermal protection function and another is transpirationcooling thermal protection function. The complex hypersonic flowfield and walltemperature are simulated and analyzed. And the simulation method is improved withpreconditioning method.The opposing jet thermal protection function is one of the important sides. In orderto know how it works, numerical simulation of hypersonic flow field with opposing jetis performed. The3D computational code to solve Navier-Stoke function has beenestablished by using AUSMPW scheme, MUSCL method, LU-SGS method and finitevolume method. The computational results are consistent with those of the referencesand the code has been validated using experimental data. By the use of the code, thecomplex hypersonic flow field with opposing jet is obtained and analyzed. To study theeffect of the intensity of opposing jet more reasonably, a new parameter has beendefined by combining the flux and the total pressure ratio. The study shows that thesame shock wave position, drag coefficient and total heat load can be obtained with thesame new parameter with different fluxes and the total pressures, and the new parameterhas qualitative relationship with the flowfield coefficients. As a base to calculate thewall temperature of the opposing-jet and inspiration leading edge, the flowfieldcharacteristic and heat load should be obtained correctly, which will support the wholethermal analysis.In the hypersonic flow field with opposing jet, there are both high speed area andlow speed area. In the low speed area, there is low speed effect, which results inconvergence deterioration and incorrect solution. However, we can introducepreconditioning method in order to accelerate the convergence of the steady solutionand obtain the numerical solution correctly. When the low speed problem is solvednumerically, the flux splitting schemes with preconditioning is deduced based on therelationship between the flux splitting schemes and system eigenvalues. In order to useflux vector splitting scheme in the preconditioning method, a new scheme is proposed based on the Roe scheme and can be used in the preconditioning method. The numericaltests of a low speed cases succeed with convergence acceleration and correct solution,which validates the preconditioning method.The plantlet transpiration cooling thermal protection function is another importantside. The physical shape of the plantlet transpiration cooling nose is built. In order toanalysis the thermal characteristic of the whole nose, both heating and cooling wayshould be considered. The methods of calculating the aerodynamic heating, such asengineering method and CFD method, are introduced. For cooling way, a distributionmodel of coolant is proposed for the later temperature calculation. With the heating andcooling data, the whole nose thermal state can be obtained with wall temperaturedistribution and coolant temperature distribution in the grooves by use of Finite VolumeElement.With the method of obtaining the thermal state of the whole nose, the transpirationcooling effect is studied. It is showed that the transpiration cooling can keep the highestwall temperature of the nose within the working temperature range of the material,which makes sure the nose keep working within the allowed temperature with seriousaerodynamic heating. The cooling effect with the plantlet cooling groove structure isdiscussed and the physical mechanism is analyzed. Combined with opposing jet thermalprotection method, the plantlet nose will work better. The synthetical thermal analysisof the opposing jet thermal protection function and the transpiration cooling thermalprotection function testifies the validity of the opposing-jet and inspiration leading edge.The experiment on flowfield around the opposing-jet and inspiration leading edgehas been conducted in the supersonic wind tunnel. The high-definitionNanoparticle-based Planar Laser Scattering (NPLS) is used to observe the flowfield.The experiment is designed based on the characteristic of the supersonic flow withopposing jet. By NPLS, the complex structure of the flowfield with opposing jet can beobserved in detail, including the bowl shock wave in front of the nose and therecompression shock wave near the reattachment of the jet layer. It is showed that theexperiment results are consistent with the calculation results, which validates thecalculation method again.