HF-plasmatron experiment and numerical simulation of heat transfer in underexpanded dissociated-nitrogen jets

Experiments on heat transfer in underexpanded supersonic jets of high-enthalpy nitrogen are performed on the VGU-4 induction high-frequency plasmatron at a pressure of 10.4 GPa in a compression chamber. At gas flow rates of 2.4 and 3.6 g/s and HF generator powers of 45 and 64 kW the heat fluxes to the copper, stainless steel, MPG-7 graphite, and quartz surfaces are measured at the stagnation point of a water-cooled cylindrical, flat-ended model, 20 mm in diameter. In the same regimes the stagnation pressures are measured. The effect of the surface catalyticity with respect to nitrogen atom recombination on the heat flux is demonstrated and the qualitative catalyticity scale of the studied materials is established. In the supersonic regimes nonequilibrium nitrogen plasma flow in the discharge channel of the plasmatron and the underexpanded jet flow past the model are numerically simulated for the experimental conditions. The experimental and calculated data on the stagnation pressures and the heat fluxes to cooled surfaces of the metals, graphite, and quartz are compared.     

克服“高超声障''的途径”

在高超声速飞行条件下, 流入冲压发动机燃烧室并降至低速的空气温度, 随着飞行马赫数增 加升得愈来愈高. 燃料与高温空气混合燃烧释放的化学能将部分转化为解离能. 这些解离能 在长度受限的尾喷管中难以充分复合形成推力, 使冲压发动机性能随飞行马赫数增大而急剧 下降. 导致冲压发动机不适应高超声速飞行器的推进要求. 将此定名为``高超声障'. 半个 世纪以来, 广泛采用``超声速燃烧'降低流入燃烧室的空气温度来克服这种障碍. 虽已取得 不少进展, 然而关键性难点仍需继续攻克. 为了多途径促进吸气推进高超声速飞行的实现, 提出克服``高超声障'的另一种思路：保持现有冲压发动机吸气与燃烧方式, 通过催化促进 燃气解离组分在尾喷管膨胀过程中的复合, 增大冲压发动机的推力, 达到满足高超声速飞行 器的推进要求.     

近空间高超声速飞行器气动特性研究的若干关键问题

在30$\sim$70km空域机动飞行的高超声速飞行器的优点是可以耦合利用所处空域的空气产生的升力和高速飞行的离心力进行远距离机动滑翔飞行,具有重要的实用价值.尽管过去数十年在高超声速流动研究方面取得显著进展,但在设计研究近空间远程滑翔的高超声速飞行器方面仍然存在许多挑战,特别是对特定飞行条件下的流动机理了解不清楚.本文介绍了作者研究团队在开展近空间高超声速飞行器有关的关键气动问题方面的研究进展,主要包括:建立了近空间高超声速飞行的流动模型,发展了系统的相关计算空气动力学方法,针对高空高速飞行条件下稀薄气体效应和真实气体效应的耦合作用影响研究了合适的滑移边界条件,考虑了不同组分存在条件下的温度、速度和压力的滑移效应影响;提出了飞行器气动外形的动态优化方法,获得了可工程实用化的高升阻比飞行器气动外形;建立了高速飞行器动稳定性理论,在实现高超声速飞行器动态稳定飞行方面取得重大进展;最后讨论了高超声速飞行器设计中进一步需要关注的若干关键技术和科学问题、可能解决的途径及其所涉及的学科发展方向.      

高超声速飞行流场中的最大氧离解度分析简

针对大气层内高超声速飞行时的化学非平衡现象,建立了沿驻点线流动的空气中氧离解度的计算模型. 模型假设氮气在氧气未充分离解时不发生离解,并且不涉及边界层内的复合反应. 理论计算发现,空气中氧的离解度随飞行高度的增加呈先增后减的非单调变化规律,其原因是由于化学反应平衡移动与非平衡效应相互作用的结果. 这一结论得到了数值模拟结果的验证,同时也解释了文献中当飞行高度较高时真实气体效应减弱的现象. 基于驻点线的近似理论模型,计算得到了轴对称钝头体绕流流场中的最大氧离解度及边界层外缘温度随飞行速度和高度变化的等值线图谱,相关结果可以为工程设计所用.     

高超声速飞行流场中的最大氧离解度分析

针对大气层内高超声速飞行时的化学非平衡现象,建立了沿驻点线流动的空气中氧离解度的计算模型. 模型假设氮气在氧气未充分离解时不发生离解,并且不涉及边界层内的复合反应. 理论计算发现,空气中氧的离解度随飞行高度的增加呈先增后减的非单调变化规律,其原因是由于化学反应平衡移动与非平衡效应相互作用的结果. 这一结论得到了数值模拟结果的验证,同时也解释了文献中当飞行高度较高时真实气体效应减弱的现象. 基于驻点线的近似理论模型,计算得到了轴对称钝头体绕流流场中的最大氧离解度及边界层外缘温度随飞行速度和高度变化的等值线图谱,相关结果可以为工程设计所用.      

Theory of an Ideal Jet Thrust Augmentor

A theory of the ideal jet thrust augmentor is presented. The conditions of optimal outflow of the active (primary) and passive (secondary) jets from the device under consideration are obtained by solving the variational problem of maximum average thrust realization. The inlet values of the mass and total enthalpy fluxes for both flows, their entropies, and the inlet value of the passive-gas momentum component parallel to the flight velocity are preassigned. These conditions correspond to the use of jet engines, including a bypass turbojet gas generator, operating in steady mode, as well as a pulsed detonation rocket engine, as the high-pressure gas source. Along with the work done by the high-pressure gas on the low-pressure one, the gases may exchange heat. The possibility of heat transfer results in an appreciable increase in ideal thrust augmentor performance.     

Attainment of maximum levels of natural convective heat transfer across major cavities using as a coolant a mixture of two pure gases instead of air

The enhancement of heat transfer in natural convection cavities is a very difficult task because of the intervening low fluid velocities. It is of fundamental and practical interest to explore alternative instruments that are power-independent and exclude surface modifications for the augmentation of heat transfer in these cavities. One feasible way for enhancing heat transfer rates passively in cavities filled with a gas is to stimulate the mechanism by natural convection of heat. The central objective of this paper is to employ a mixture of two pure gases that yields levels of heat transfer increments that are unattainable by each pure gas acting along (or even by air). In general, dimensional analysis insinuates that four transport properties affect natural convection flows: density, isobaric specific heat capacity, dynamic viscosity and thermal conductivity. Simple correlation equations of power form are useful to engineers for a quick estimate of the magnitudes of the space-mean heat transfer coefficient. Detailed computations were made for four different gases: air, pure helium, pure argon, and a mixture of pure helium and pure argon and the relative merits of each of them have been discussed. Five major cavities of relevance in applications of thermal engineering have been analyzed in this work. Received on 6 August 1999     

高超声速激波风洞研究进展

回顾了高超声速激波风洞的研制与发展,并依据高超声速实验研究对地面实验模拟技术的要求,分别介绍了应用轻气体、自由活塞和爆轰驱动技术研制的主要激波风洞的性能、特点和存在问题.重点介绍了爆轰驱动高焓激波风洞的3种主要运行模式:反向、正向爆轰驱动与双爆轰驱动. 根据这些运行模式的工作原理,分析了应用这些驱动技术产生的高温、高压气源的特点,探讨了不同驱动技术可能影响激波风洞性能的关键问题与解决方法.目前发展的激波风洞已经能够用于开展马赫数3$\sim$30的高超声速流动的试验模拟研究,但是试验气流的品质还不能满足高超声速科技研究的需求.为了获得可靠的实验结果, 通过不断改进、完善、提高激波风洞的性能,尽可能复现高超声速飞行条件是今后主要的研究方向.      

Generalized reference enthalpy formulations and simulation of viscous effects in hypersonic flow

A generalized reference enthalpy formulation for the skin friction, heat transfer and radiation-equilibrium temperature distributions over aerodynamic surfaces in attached hypersonic / hyperenthalpic flow is proposed. The formulation, which has been extensively employed in various forms by numerous investigators in the perfect gas regime, has also been recently demonstrated to provide adequate estimates of the heat transfer distribution in thermochemically active high enthalpy flow conditions when coupled to thermochemically active Euler solutions. It is now used to reveal the relevant similitude parameters for viscous effects in hypersonic flow, and the importance of the temperature distribution across the boundary layer and of the temperature-viscosity relation. It is shown that, although reproduction of the flight total flow enthalpy as well as surface temperature is the obvious solution for full viscous simulation in (perfect gas) hypersonic flow, the hot surface testing requirement and, in a number of practical applications, also the hot flow requirement may be relaxed with reasonably small error that can be of the same order as the measurement accuracy in present-day hypersonic testing. This similitude error, however, may increase significantly in cases exhibiting strong viscous/inviscid interaction or when the laminar-turbulent transition process becomes important. In this respect, alternative full simulation solutions, which are less demanding in terms of reproduction of the high levels of flight freestream and surface temperature or even Reynolds number, are discussed. Received 6 May 1997 / Accepted 8 October 1997     

Influence of real gas properties on the integral aerodynamic coefficients

The results of mathematically modeling axisymmetric hypersonic flow past an ellipsoid are presented. The calculation data are obtained on the basis of a numerical solution of the complete Navier—Stokes equations using a finite-difference method. The investigation is carried out for the problem of laminar flow over the windward side of the body. The effect of the elongation of the ellipsoid on the total heat flux and the viscous and pressure drag is considered. Results are obtained for three different gas (air) models: the perfect gas, chemical equilibrium and chemical nonequilibrium models. For the latter model various sets of catalyticity coefficients are considered. The effect of the real properties of air on the integral aerodynamic characteristics of ellipsoids with different elongations is analyzed.Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No.4, pp. 156–164, July–August, 1992.     

热喷干扰气体模型对飞行器气动特性影响分析

针对不同气体模型对高超声速飞行器喷流反作用控制系统(RCS)热喷干扰流场模拟的计算效率和准确性问题, 基于喷流燃气物理化学模型, 通过数值求解含化学反应源项的三维N-S方程, 建立了飞行器RCS热喷干扰流场数值模拟方法, 分别采用化学反应流、反应冻结流、二元异质流以及空气喷流四种气体模型开展了典型外形热喷干扰流场的数值模拟, 研究了不同气体模型对热喷干扰流场结构、飞行器气动力热特性的影响, 分析了不同马赫数、飞行高度下的变化规律. 研究表明: 化学反应流模型计算精度较高, 计算与风洞试验数据的吻合程度优于其他三种简化模型; 在本文的低空条件下, 采用简化模型进行热喷干扰流场数值模拟, 会低估分离区大小, 使飞行器气动力特性预测出现偏差, 同时也会低估表面热环境, 对防热系统设计不利, 随着马赫数增加, 简化模型对气动力热特性预估的误差进一步增大, 同时不同简化模型之间的差异也进一步增大; 飞行高度较高时, 模型之间的差异减小, 此时可采用简化模型进行计算以提高计算效率. 本文的研究结果可为飞行器热喷干扰流场数值模拟及喷流反作用控制系统设计提供参考.      

An implicit upwind parabolized Navier-Stokes code for chemically nonequilibrium flows

The previously developed single-sweep parabolized Navier-Stokes (SSPNS) space marching code for ideal gas flows has been extended to compute chemically nonequilibrium flows. In the code, the strongly coupled set of gas dynamics, species conservation, and turbulence equations is integrated with the implicit lower-upper symmetric Gauss-Seidel (LU-SGS) method in the streamwise direction in a space marching manner. The AUSMPW+ scheme is used to calculate the inviscid fluxes in the crossflow direction, while the conventional central scheme for the viscous fluxes. The k-g two-equation turbulence model is used. The revised SSPNS code is validated by computing the Burrows-Kurkov non-premixed H2/air supersonic combustion flows, premixed H2/air hypersonic combustion flows in a three-dimensional duct with a 15° compression ramp, as well as the hypersonic laminar chemically nonequilibrium air flows around two 10° half-angle cones. The results of these calculations are in good agreement with those of experiments, NASA UPS or Prabhu's PNS codes. It can be concluded that the SSPNS code is highly efficient for steady supersonic/ hypersonic chemically reaction flows when there is no large streamwise separation.     

Body shape with minimum total radiant energy flux toward its surface

During a space vehicle's entry into a planet's atmosphere at hypersonic speed one of the important problems is the aerodynamical surface heating due to convective and radiant heat fluxes from the gas after passing through a strong shock wave. Due to the high destructive action of this heating, an important problem is the selection of the aerodynamic shape allowing the minimum heat influx to its surface. The problem of determining the shapes of an axisymmetric body from the condition of minimum total convective heat flux along the lateral face of the body was considered under various assumptions in [1–7]. There are a number of entry conditions (for example, into the earth's atmosphere with a speed of 11 km/ sec at an altitude of about 60 km [12]) during which the radiative component becomes dominant in the total heat flux toward the body. A numerical solution of the problem of hypersonic flow of a nonviscous, non-heat-conducting radiating gas around a body is obtained at this time only for a limited class of bodies and primarily for certain entry conditions (for example, [8–12]). On the basis of these calculations it is impossible to make general conclusions concerning arbitrary body shapes. Therefore, approximate methods were proposed which permit the distribution of radiant heat flux to be obtained for an arbitrary axisymmetric body in a wide range of flight conditions [13–15]. In the present work an expression is derived for the total radiant heat flux over the entire body surface and similarity criteria are found. A variational problem is formulated to determine the shape of an axisymmetric body from the condition of minimum total radiant-heat flux over the entire body surface. It is solved analytically for the class of thin bodies and in the case of a strongly radiating gas.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 4, pp. 84–89, July–August, 1976.     

Supersonic low-concentration dusty-gas flow past a plane cylinder in the presence of an oblique shock wave interacting with the bow shock

The motion of an inertial dispersed admixture near a plane cylinder immersed in a steady-state hypersonic dusty flow in the presence of an oblique shock wave interacting with the bow shock is considered. It is assumed that the free-stream particle mass concentration is small and the particles do not affect the carrier flow. The III and IV shock wave interaction regimes are considered. The gas flow parameters in the shock layer are calculated from the numerical solution of the full Navier-Stokes equations for the perfect gas. A TVD second-order finite-difference scheme constructed on the basis of a finite volume method is used. For calculating the dispersed-phase parameters, including the concentration, the full Lagrangian method is used. On a wide range of variation of the particle inertia parameters, the patterns of the particle trajectories, velocity, concentration, and temperature in the shock layer are studied. The possibility of aerodynamic focusing of the particles behind the shock wave intersection point and the formation of narrow beams with a high particle concentration is revealed. These beams impinge on the cylinder surface and result in a sharp increase in the local heat fluxes. The maximal possible increase in the heat fluxes caused by the particles colliding with the cylinder surface is estimated for the flows with and without the incident oblique shock wave.     

The effects of variable specific heat on the stability of hypersonic boundary layer on a flat plate

The gas temperature within hypersonic boundary layer flow is so high that the specific heat of gas is no longer a constant but relates to temperature. How variable specific heat influences on boundary layer flow stability is worth researching. The effect of the variable specific heat on the stability of hypersonic boundary layer flows is studied and compared with the case of constant specific heat based on the linear stability theory. It is found that the variable specific heat indeed has some effects on the neutral curves of both the first-mode and the second-mode waves and on the maximum rate of growth also. Therefore, the relationship between specific heat and temperature should be considered in the study of the stability of the boundary layer.     

高超声速喷管设计理论与方法

在高超声速飞行技术领域,特别是涉及到高焓气体流动的研究,高超声速风洞试验仍然是目前最可靠的研究手段.风洞流场的品质是高超声速风洞研发最重要的一项性能指标,其取决于喷管设计采用的理论与方法,也是风洞设计最关注的一项核心技术.针对二维轴对称型面喷管设计,本文首先综述了传统高超声速喷管设计的主要理论和常用方法,它们在高超声速...     

Effect of a magnetic field on the flow in the vicinity of the stagnation point of a blunt body with ablation of a protective layer

During hypersonic flow around a blunt-nosed body, the gas which passes through the bow shock is heated to high temperatures, where dissociation, ionization, and inverse phenomena (recombination) take place in the gas. If an ionized gas moves in a magnetic field, the ponderomotive force which is set up changes the nature of its motion close to the stagnation point, decreasing the frictional stress and heat transfer at the wall (at the contact surface of the gas and the body about which the gas flows). In this case, the intense heat fluxes from the strongly heated gas to the body about which the gas flows cause phase changes in the surface of the body (melting, sublimation, etc.). These processes, in turn, affect the flow in the vicinity of the stagnation point due to realization of the heat of phase transition, the conduction of heat from the entrained mass, and the diffusion of evaporating material into the boundary layer. References [1, 2] are devoted to a study of the joint influence of the magneto-gasdynamic and ablation effects. The magnetogasdynamic layers and the wall profile of the external velocity (flow around wedges) are discussed in [1], and special cases of such boundary layers-flow close to the stagnation line (the two-dimensional case) and close to the stagnation point (the axisymmetric case) of a blunt body are considered in [2]. Melting and evaporation are taken into account by setting the longitudinal and the transverse velocity components at the wall not equal to zero-the first taking into account the flow of the molten material and the second pyrolysis of the vapor of the surface material into the gaseous boundary layer. However, the values of these components, also the enthalpy on the wall h_w(in[1, 2] h_W 0), are not known beforehand and must be determined from the boundary conditions at the wall which express the mass and heat balances. The general formulation of the problem given in the gasdynamics case by G. A. Triskii in [3, 4], and elsewhere includes a consideration of the boundary-layer equations in the gas, the boundary-layer equations in the melted zone, and the heat conductivity equations in the solid with boundary conditions at the outer edge of the boundary layer, on the gas-molten zone interface, on the molten zone-solid interface, and inside the solid. This approach to the problem can also be utilized in the magnetogasdynamic case, as it is in this article with certain simplifying assumptions as compared with [3, 4]. In this sense, the present article is an extension of the results of [3, 4] to the field of magnetogasdynamics.In conclusion, the author thanks K. A. Lur'e for proposing the subject and useful discussions.     

Three-dimensional supersonic flow past blunt bodies with contour discontinuities,with account for equilibrium and frozen state of the gas in the shock layer

An algorithm is devised for calculating by the finite difference method the supersonic flow region for three-dimensional steady-state flow of a viscous gas past a blunted body with many contour discontinuities. The state of this gas at high hypersonic flight speeds can be characterized by equilibrium or frozen physicochemical processes. Generally speaking, any arbitrary number and sequence of either compression or expansion discontinuities is permitted. The computational scheme adopted provides identification of the vortex layers, regions with different equations of state, and so on. We use a flow model that is either frozen throughout the entire shock layer or only in the portion of the layer adjacent to the body surface. The pressure at certain points on the surface of spherically blunted cones with half-angles θ?10° may differ by a factor of 2 or more in equilibrium and frozen flows. Example calculations are presented, and the results are analyzed.     

高超声速飞行复现风洞理论与方法

针对高超声速飞行伴随的热化学反应流动,本文回顾了郭永怀先生的科研理念和学科布局,综述了他亲手成立的高温气动团队在高超声速飞行风洞实验模拟理论与方法方面的研究进展.高温气体的迅速产生与迅速应用是一种理想的风洞运行方法,而激波管就是这样一种实验装备.论文首先介绍了激波管技术的基本理论与方程,指出将其用于高超声速流动实验模拟时所具有的独特优势.然后讨论了应用激波风洞复现需要的高超声速飞行状态的可行性、基本方程和需要解决的关键问题.针对这些关键问题,进一步介绍了如何应用爆轰现象研发激波风洞驱动技术的理论,并给出了基于爆轰驱动方法的技术发展和工程应用验证.最后,论文介绍了爆轰驱动激波风洞的界面匹配条件,该条件奠定了长实验时间激波风洞运行基础,是其他驱动方法尝试解决而没能完全解决的难题.高温气动团队关于高超声速飞行复现风洞的理论与技术研究,实现了郭永怀先生的战略规划,成就了国际领先的高超声速热化学反应流动研究平台.