Mechanism of unsteady aerodynamic heating with sudden change in surface temperature

The characteristics and mechanism of unsteady aerodynamic heating of a transient hypersonic boundary layer caused by a sudden change in surface temperature are studied. The complete time history of wall heat flux is presented with both analytical and numerical approaches. With the analytical method, the unsteady compressible boundary layer equation is solved. In the neighborhood of the initial and final steady states, the transient responses can be expressed with a steady-state solution plus a perturbation series. By combining these two solutions, a complete solution in the entire time domain is achieved. In the region in which the analytical approach is applicable, numerical results are in good agreement with the analytical results, showing reliability of the methods. The result shows two distinct features of the unsteady response. In a short period just after a sudden increase in the wall temperature, the direction of the wall heat flux is reverted, and a new inflexion near the wall occurs in the profile of the thermal boundary layer. This is a typical unsteady characteristic. However, these unsteady responses only exist in a very short period in hypersonic flows, meaning that, in a long-term aerodynamic heating process considering only unsteady surface temperature, the unsteady characteristics of the flow can be ignored, and the traditional quasi-steady aerodynamic heating prediction methods are still valid.

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.

AERODYNAMIC HEATING IN THE REGION OF SHOCK AND TURBULENT BOUNDARY LAYER INTERACTION INDUCED BY A CYLINDER

Detailed distributions of heat flux in the region of shock wave and turbulent boundary layer interaction induced by a cylinder were measured in the shock tunnel Oil flow patterns and Schlieren photo- graphs were taken.Empirical relations were given for determining separation shock angle,peaks of heat flux and their locations on both cylinder leading edge and flat plate surface,and other characteristic parameters of the interaction region.

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.

Comparisons and applications of numerical simulation methods for predicting aerodynamic heating around complex configurations

Numerical simulation methods of aerodynamic heating were compared by considering the inuence of numerical schemes and turbulence models,and attempting to investigate the applicability of numerical simulation methods on predicting heat flux in engineering applications. For some typical cases provided with detailed experimental data,four spatial schemes and four turbulence models were adopted to calculate surface heat flux. By analyzing and comparing,some inuencing regularities of numerical schemes and turbulence models on calculating heat flux had been acquired. It is clear that AUSM＋-up scheme with rapid compressibilitymodified high Reynolds number k鈥撓？model should be appropriate for calculating heat flux. The numerical methods selected as preference above were applied to calculate the heat flux of a 3-D complex geometry in high speed turbulent flows. The results indicated that numerical simulation can capture the complex flow phenomena and reveal the mechanism of aerodynamic heating. Especially,the numerical result of the heat flux at the stagnation point of the wedge was well in agreement with the prediction of Kemp鈥揜iddel formula,and the surface heat flux distribution was consistent with experiment results,which implied that numerical simulation can be introduced to predict heat flux in engineering applications.

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.

Comparative study on aerodynamic heating under perfect and nonequilibrium hypersonic flows

In this study, comparative heat flux measurements for a sharp cone model were conducted by utilizing a high enthalpy shock tunnel JF-10 and a large-scale shock tunnel JF-12, responsible for providing nonequilibrium and perfect gas flows, respectively. Experiments were performed at the Key Laboratory of High Temperature Gas Dynamics(LHD), Institute of Mechanics, Chinese Academy of Sciences. Corresponding numerical simulations were also conducted in effort to better understand the phenomena accompanying in these experiments. By assessing the consistency and accuracy of all the data gathered during this study, a detailed comparison of sharp cone heat transfer under a totally different kind of freestream conditions was build and analyzed. One specific parameter, defined as the product of the Stanton number and the square root of the Reynold number, was found to be more characteristic for the aerodynamic heating phenomena encountered in hypersonic flight. Adequate use of said parameter practically eliminates the variability caused by the deferent flow conditions, regardless of whether the flow is in dissociation or the boundary condition is catalytic. Essentially, the parameter identified in this study reduces the amount of ground experimental data necessary and eases data extrapolation to flight.

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.

Thermal-mechanical response of microscale functional film for infrared window

Infrared window in hypersonic missile usually suffers complex aerodynamic force/heat during high-speed flight.A finite element method was adopted to simulate the thermal and stress response of microscale functional film for infrared window under different aerodynamic heats/forces conditions.Temperature and stress distribution were obtained with different heat fluxes.There is almost constant stress distribution along the film thickness except a sudden decrease near the substrate.The maximum stresses are located at the points which are 0.3 mm away from the edges.Different film materials result in different stress values.The temperature and stress in ZrN are larger than those in Y2O3.Besides the numerical simulation,an oxygen propane flame jet impingement test was performed to investigate thermal shock failure of the infrared window.Some place of the window surface has spots damage and some place has line crack damage after thermal shock.

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.

ADAPTIVE FINITE ELEMENT METHOD FOR HIGH-SPEED FLOW-STRUCTURE INTERACTION

An adaptive finite element method for high-speed flow-structure interaction is pre- sented.The cell-centered finite element method is combined with an adaptive meshing technique to solve the Navier-Stokes equations for high-speed compressible flow behavior.The energy equation and the quasi-static structural equations for aerodynamically heated structures are solved by applying the Galerkin finite element method.The finite element formulation and computational procedure are de- scribed.Interactions between the high-speed flow,structural heat transfer,and deformation are studied by two applications of Mach 10 flow over an inclined plate,and Mach 4 flow in a channel.

HP nonlinear guidance law design based on smooth sliding mode control

To eliminate the perturbation of interceptor detection induced by aerodynamic heating,the head pursuit (HP) guidance law for three-dimensional interception was presented. The guidance law positioned the interceptor ahead of the target on its flight trajectory,and the speed of interceptor was required to be lower than that of the target. On the basis of a novel head pursuit three-dimensional guidance model,a nonlinear guidance law was developed based on smooth sliding mode control theory. At the same time,a special observer was designed to estimate the target acceleration,and a numerical example on maneuvering ballistic target interception verified the effectiveness of the presented guidance law.

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.

Investigation of thermal protection system by forward-facing cavity and opposing jet combinatorial configuration

This paper focuses on the usage of the forward-facing cavity and opposing jet combinatorial configuration as the thermal protection system （TPS） for hypersonic vehicles. A hemispherecone nose-tip with the combinatorial configuration is investigated numerically in hypersonic free stream. Some numerical results are validated by experiments. The flow field parameters, aerodynamic force and surface heat flux distribution are obtained. The influence of the opposing jet stagnation pressure on cooling efficiency of the combinatorial TPS is discussed. The detailed numerical results show that the aerodynamic heating is reduced remarkably by the combinatorial system. The recirculation region plays a pivotal role for the reduction of heat flux. The larger the stagnation pressure of opposing jet is, the more the heating reduction is. This kind of combinatorial system is suitable to be the TPS for the high-speed vehicles which need long-range and long time flight.

Thermal Protection Efficiency of Forward-facing Cavity and Opposing Jet Combinational Configuration

To deal with the thermal protection of high speed vehicle, the cooling efficiency of a combinatorial thermal protection configuration which is composed of the forward-facing cavity and opposing jet is investigated. The numerical simulation result is validated by experiment and the flow field parameters, aerodynamic force and heat flux distribution are obtained. The detailed numerical simulation results show that this kind of combinatorial thermal protection configuration has an excellent effect on cooling the surface of the nosetip. By adding of the opposing jet with a small total pressure, it can avoid the disadvantage to the control performance of the aircraft which is caused by the cavity oscillating flow. And the low stagnation pressure is propitious to simplify the opposing jet system. The location of the recirculation region has a significant impact of the aerodynamic heating. The heat flux along outer body surface of the nosetip does not increase with the stagnation pressure of opposing jet decreases monotonically.

Numerical analysis on the heat transfer of three types of nozzles for the hypersonic long-run wind tunnel

The hypersonic long-run scramjet test tunnel is one of the key ground facilities for the studies of ramjet/scramjet and hypersonic thermal management.Due to the significantly large heat loading,the nozzle of the tunnel facility demands effective cooling protection.In this work,the two-dimensional,three-dimensional and axisymmetric Mach 6.5 nozzles at an inlet total temperature of 1840 K and a total pressure of 6.4 MPa were studied with main focuses on the properties of aerodynamic heating of nozzles.The present work aims to provide insights into the design of an effective cooling system for the nozzle and other components of the hypersonic long-run wind tunnel.

Numerical investigation on the thermal protection mechanism for blunt body with forward-facing cavity

Numerical experiments are carried out using the standard hypersonic ballistic-type model(HB-2) to investigate the effect of forward-facing cavity on the aerodynamic heating. A general concept is proposed which utilizes the flow disturbances generated passively in the nosed subsonic region to weaken the detached shock wave. Several aspects are mainly studied, including shock shape and standoff distance, surface heat flux and pressure, flowfield feature and cooling mechanism. The numerical results indicate that shock strength and standoff distance increase with an increase in the L/D ratio of the cavity. Interestingly, a bulge structure of the detached shock associated with a deep cavity is observed for the first time. Local surface heat flux and pressure around the concave nose are much lower respectively than those at the stagnation point of the baseline model. In addition, both surface heat and pressure reductions are proportional to the L/D ratio. A negative heating phenomenon may occur in the vicinity of a sharp lip or on the base wall of a deep cavity. If the L/D ratio exceeds 0.7, the detached shock appears as a self-sustained oscillation which can be referred to as the cooling mechanism.

Research on Cooling Effect with Coolant Groove Structure Parameters in Platelet Transpiration Cooled Nosetip

A platelet transpiration cooled nosetip is considered as thermal protection system （TPS） to prevent hypersonic ve- hicle from the serious aerodynamic heating. Based on the one dimensional flow model, a distribution model of coolant is proposed for the temperature calculation. When Si = Sj （i, j=1,… 24）, the first cooling effect parameter Pmax is proposed and its relationship with total mass flux and Sc0/Si is investigated. The result shows that Pmax increases while the total mass flux increases, and when the mass flux is fixed, Pmax increases rapidly at the be- ginning and then turns to a nearly stable value while Sc0/Si increases. Then under the precondition of cooling ef- fect, we fix Sc0/Si to insure there is enough space for the pipe. Numerical investigation shows the design of the nosetip makes the transpiration cooling extremely effective. In order to reduce the temperature difference on the nosetip, the second cooling effect parameter Pdiff is proposed and different Pdiff with different 0gi （i=1,..., 23） are analyzed. According to the cases we design, Pdiff decreases while the upstream 0gi decreases or the down- stream 0gi increases. The best result among cases shows Pdiff is reduced by 15.1%.