Investigations On The Application Feasibility Of Transpiration Cooling On Hypersonic Vehicles

With the increase of hypersonic vehicle speed,worse thermal environment occurs over the critical structures like the leading edges of nose and wing,or the thrust chamber of the engine.Transpiration cooling,as the most efficient active thermal protection technique,can not only increase the cooling efficiency and improve the temperature uniformity,but also supply a certain amount of coolant to a certain high temperature region at a certain time actively,hence the cooling process is controllable,which gains strength to expand the cooling area,increase the reusability and realize intelligent control of the thermal protection system.However,many concrete issues should be discussed and solved to promote its practical application.For example,the aerodynamic heat and force are extremely non-uniform on the hypersonic leading edge,the transpiration cooling system should supply a large portion of coolant to the stagnation region to resist the tremendous thermal loads,simultaneously,have to overcome the obstacle of high aerodynamic force at the same location.It's tough to realize a reasonable coolant allocation matching with the external heat and force distribution.In this dissertation,an entire-field-coupled numerical approach is promoted and validated to investigate the cooling mechanisms and non-uniform coolant supply strategies of transpiration cooling.The main work and conclusions are as follows:(1)A discrete transpiration cooling structure is designed on a flat plate to realize a reasonable coolant allocation by adjusting the porosity of the separated porous matrixes.The numerical results indicate that:The heat transfer on the coupled wall plays a more important role when the coolant injection rate increases.When compared with single transpiration cooling,the heat transfer on the coupled wall has more significant influence on discrete transpiration cooling.Besides,a porosity prediction model is established to calculate all kinds of porosity combination that is able to realize the same coolant allocation.The model is helpful to further improve the discrete transpiration cooling system to meet additional constraints related to weight and mechanic strength.(2)To deal with the extremely non-uniform aerodynamic heat and force on the hypersonic leading edges,the application feasibility of the aforementioned discrete transpiration cooling is discussed.Firstly,single transpiration cooling structures are used to observe the influence of coolant injection rate and porous matrix length on the downstream film cooling effects derived from the upstream transpiration cooling.It's found that single transpiration cooling is unable to protect a hypersonic leading edge with reasonable coolant consumption.Second,the cooling effects of single and discrete transpiration cooling are compared,it's demonstrated that by adjusting the coolant allocation into different porous matrixes,discrete transpiration cooling is able to expand the efficient thermal protection area with the less coolant consumption.At last,the entire-field-coupled numerical approach is adopted to check the feasibility of adjusting the coolant allocation by changing the porosity of different porous matrixes.A certain porosity combination is selected out rapidly to realize the target coolant allocation.(3)A double layer combined cooling structure is proposed on a flat plate with an inner film cooling layer and an outer transpiration cooling layer.The numerical results indicate that:The outer transpiration cooling layer can effectively restrain the lift-off effect of the coolant ejecting out of the film holes.The cooling effects improve as the thickness of the porous layer increases,but deteriorate as the porosity rises.Besides,the cooling performances of uniform film-hole layout is compared with that of three non-uniform film-hole layouts.It's found that the cooling uniformity can be improved significantly by adjusting the film-hole layout.(4)Considering the thermal protection issue of hypersonic leading edges,the aforementioned double layer combined cooling is applied under hypersonic conditions.The research firstly analyzes the influences of film-hole layout on the comprehensive cooling effects,and it's concluded that as the number of film-hole increases,the cooling performances improve first and deteriorate then,hence,a best design exists.After that,the cooling mechanisms in different regions are discussed,and it's found that a low temperature area like an umbrella forms over the stagnation region when coolant ejects out of the through hole,the coolant diffuses from the other film holes into the porous matrix and provides transpiration cooling effects in the middle region,and the convective cooling effects along the inner wall of the coolant channel protect the downstream region.Finally,as a peak temperature still occurs near the stagnation region,a new strategy which locally adjusts the hole-depth into the porous matrix is exhibited and numerical simulations are conducted to check its feasibility.