CALCULATING METHOD OF AERODYNAMIC HEATING FOR HYPERSONIC AIRCRAFTS

A new calculating method of aerodynamic heating for unsteady hypersonic aircrafts with complex configuration is presented.This method,which considers the effects of high temperature chemical non-equilibrium and the heat transfer process in thermal protection structure,is based on the combination of the inviscid outerflow solution and the engineering method,where the Euler solver provides the flow parameters on boundary layer edge for engineering method in aerodynamic heating calculation.A high efficient interpolation technique,which can be applied to the fast computation of longtime aerodynamic heating for hypersonic aircraft,is developed for flying trajectory.In this paper,three hypersonic test cases are calculated,and the heat flux and temperature distribution of thermo-protection system are shown.The numerical results show the high efficiency of the developed method and the validation of thermal characteristics analysis on hypersonic aerodynamic heating.

Effects of Wall Emissivity on Aerodynamic Heating in Scramjets

The effects of the wall emissivity on aerodynamic heating in a scramjet are analyzed.The supersonic turbulent combustion flow including radiation is solved in the framework of a decoupled strategy where the flow field is determined first and the radiation field next.In particular,a finite difference method is used for solving the flow while a DOM(iscrete ordinates method)approach combined with a WSGGM(weighted sum of gray gases)model is implemented for radiative transfer.Supersonic nonreactive turbulent channel flows are examined for a DLR hydrogen fueled scramjet changing parametrically the wall emissivity.The results indicate that the wall radiative heating rises greatly with increasing the wall emissivity.As the wall emissivity rises,the radiative source and total absorption increase,while the incident radiation decreases apparently.Notably,although the radiative heating can reach a significant level,its contribution to the total aerodynamic heating is relatively limited.

Numerical study on the influence of initial ambient temperature on the aerodynamic heating in the tube train system

The evacuated tube transportation has great potential in the future because of its advantages of energy saving and environmental protection.The train runs in the closed tube at ultra-high speed.The heat quantity generated by aerodynamic heating is not easy to spread to external environment and then accumulates in the tube,inducing the ambient temperature in the tube to rise gradually.In this paper,a three-dimensional geometric model and the Shear Stress Transport(SST)κ-ωturbulence model are used to study the influence of initial ambient temperature on the structure of the flow field in the tube.Simulation results show that when the train runs at transonic speed,the supersonic flow region with low temperature and low-pressure is produced in the wake.The structure of the flow field of the wake will change with the initial ambient temperature.And the higher the initial ambient temperature is,the shorter the low temperature region in the wake will be.The larger temperature difference caused by the low temperature region may increase the temperature stress of the tube and affect the equipment inside the tube.Consequently,the temperature inside the tube can be maintained at a reasonable value to reduce the influence of the low temperature region in the wake on the system.

ILES of an array of three subsonic counter-flow jets issuing from a wing leading edge exposed to hypersonic aerodynamic heating

The flow of an active thermal protection system exploiting subsonic counter-flow jets for wing leading edges of hypersonic vehicles is numerically studied on the basis of the three dimensional Navier-Stokes equations.The coolant air issuing from around the stagnation point as an array of three jets spreads over both the upper and the lower sides of the cylinder surface and about 40~60%cooling effectiveness is achieved in the range up to 5 degrees angle of attack despite the occurrence of various three-dimensional fluid-dynamic instabilities.The numerical scheme is second order accurate but simple inclusion of high order polynomial approximation in the reconstruction enables the capturing of finer structure of the flow field.

Prediction of Aerothermal Environment and Heat Transfer for Hypersonic Vehicles with Different Aerodynamic Shapes Based on C++

This research paper discusses constructing a unified framework to develop a full-rate scheme for hypersonic heating calculations. The method uses a flow tracing technique with normal phase vector adjustment in a non-structured delineated grid combined with empirical formulations for convective heat transfer standing and non-standing heat flow engineering. This is done using dev-C++ programming in the C++ language environment. Comparisons of the aerodynamic thermal environment with wind tunnel experimental data for the Space Shuttle and Apollo return capsules and standing point heat transfer measurements for the Fire II return capsule was carried out in the hypersonic Mach number range of 6 - 35 Ma. The tests were carried out on an 11th Gen Intel(R) Core(TM) i5-1135G7 processor with a valuable test time of 45 mins. The agreement is good, but due to the complexity of the space shuttle tail, the measurements are still subject to large errors compared to wind tunnel experiments. A comparison of the measured Fire-II return capsule standing-point heat values with the theory for calculating standing-point heat fluxes simulated using Fay & Riddell and wind tunnel experiments is provided to verify the validity of this procedure for hypersonic vehicle heat transfer prediction. The heat fluxes assessed using this method for different aerodynamic profiles of hypersonic vehicles agree very well with the theoretical solution.

Numerical Study of High-Temperature Nonequilibrium Flow around Reentry Vehicle Coupled with Thermal Radiation

Accurate aerodynamic heating prediction is of great significance to current manned space flight and deep space exploration missions.The temperature in the shock layer surrounding the reentry vehicle can reach up to 10,000 K and result in remarkable thermochemical nonequilibrium,as well as considerable radiative heat transfer.In general,high-temperature flow simulations coupled with thermal radiation require appropriate numerical schemes and physical models.In this paper,the equations governing hypersonic nonequilibrium flow,based on a three-temperature model combined with a thermal radiation solving approach,are used to investigate the radiation effects in the reentry shock layer.An axisymmetric spherical case shows that coupling the flow-field simulation with radiation has a scarce influence on the convective heating prediction,but has some impact on the radiative heating calculation.In particular,for the Apollo capsule reentry,both the absorption coefficient and incident radiation are remarkable inside the shock layer.The radiative heating maximum reaches nearly 38%of that of the convective heating making a considerable contribution to the total aerodynamic heating.These results indicate that in the hypersonic regime,in order to account for the total heating,it is necessary to simulate the high-temperature thermochemical nonequilibrium flows coupled with thermal radiation.

Transitional wave configurations between Type Ⅲ and Type Ⅳ oblique-shock/bow-shock interactions

The interactions of oblique/bow shock waves are the key flow phenomena restricting the design and aerothermodynamic performance of high-speed vehicles.Type Ⅲ and Type Ⅳ Shock/Shock Interactions(SSIs)have been extensively investigated,as such interactions can induce abnormal aerodynamic heating problems in hypersonic flows of vehicles.The transition process between these two distinct types of shock/shock interactions remains unclear.In the present study,a subclass of shock/shock interaction configuration is revealed and defined as Type Ⅲa.Type Ⅲa interaction can induce much more severe aerodynamic heating than a Type Ⅳ interaction which was ever reported to be the most serious in literature.The intense aerodynamic heating observed in this configuration highlights a new design point for the thermal protection system of hypersonic vehicles.A secondary Mach interaction between shock waves in the supersonic flow path of a Type Ⅲ configuration is demonstrated to be the primary mechanism for such a subclass of shock/shock interaction configuration.

Influence of Metal Material Properties on Heat and Mass Transfer into Thermal Protection Surface with Phenomenological Catalytic Model

Surface heterogeneous catalysis in a high-enthalpy dissociated environment leads to a remarkable enhancement of aerodynamic heating into the thermal protection surface of hypersonic aircraft.To more accurately predict this catalytic heating,a kinetic catalytic model was constructed.This model involved four elementary reactions,the rates of which were determined on mean-field approximation and surface steady-state reaction assumption.By coupling this model into the viscous wall boundary condition of computational fluid dynamics(CFD)solver,the influences of metal material catalytic properties on heat and mass transfer into thermal protection materials were numerically investigated.Numerical results showed that atomic oxygen recombination catalyzed by surface material accounts for a major contribution to aerodynamic heating and thus variation in recombination rates from different materials leads to the significant difference in surface heat fluxes.From a comparative analysis of various materials,the catalytic activity increases from the inert platinum(Pt)to nickel(Ni)and finally to the active copper(Cu).As a result,the catalytic heating on Cu surface was more than twice of that on Pt surface.Further parametrical research revealed that the proper layout of inert material at the nose of aircraft could prevent stagnation catalytic heating from thermal damage by carrying near-wall dissociated atoms from the stagnation zone downstream.The material-relied heterogeneous catalysis mechanism in this study provides some technical support for the thermal protection system design of hypersonic aircraft.