流体式高超声速可调进气道流动机理及工作特性分析

针对基于二次流控制的定几何高超声速可调进气道设计概念,给出了其具体的流道实现方案,而后通过全流道仿真分析,检验了该可调进气道在马赫数4~6范围内的可实现性,获得了其工作特性,并对弯曲激波后的总压损失特性、二次流的能量获取及消耗机制等流动机理进行了专门分析。结果表明：该流体式可调进气道能够依靠自身高压驱动二次流来实现对口部波系的调节,使进气道在低马赫数下的流量系数相对于常规定几何高超声速进气道提高24%以上,总压恢复提高7%左右,且最大二次流消耗量只占了进气道捕获流量的1.6%左右。另外,虽然弯曲激波的波后总压和马赫数分布表现出了一定的不均匀性,但是其平均总压恢复系数与相同倾角平面激波相比下降不大。二次流循环流动所消耗的机械能由外部外流剪切力做功补充,而二次流注入会使当地边界层的速度型变得瘦弱,形状因子增大。     

Uniform solution for the flow past a slender body with an attached shock wave in radiative magnetogasdynamics

A systematic perturbation scheme is used to study the propagation of a weak shock wave attached to a slender body in a supersonic flow of plasma with thermal radiation and investigate as to how the coupling between the radiative transfer and magneto-hydrodynamic phenomena affects the flow field. The analytical solution of the flow field has been presented up to the second order of ε. The shape of the shock wave attached to the slender body of revolution is obtained, which however can be expressed explicitly in terms of known functions when the radiative decay length is of the same order as the typical body length. Also, the shock angle at the tip of the projectile is obtained.     

一种正向乘波前体设计方法初步研究

以定楔角乘波体设计方法为基础,研究了影响高超/超声速乘波体"乘波"的主要因素,给出了前体前缘实际气流压缩角的确定方法及影响因素,可知在相同的来流马赫数和压缩角δ下,随着前缘角θ和气流与前缘夹角α的增加,实际气流偏转角γ减小。据此,基于幂函数进气道前体构形,给出了前缘激波不脱体的限制条件及具体的判定方法,分析了乘波体典型几何特征参数对前缘激波不脱体的影响规律,结果显示在相同的来流马赫数和压缩角度下,增大前缘形状因子n,减小前体的长宽比L/W及增大前缘角均有利于激波不脱体。根据给出的前体几何参数对前缘激波脱体的影响规律曲线,对一种"前体几何外形构造+前缘激波附体条件限制"的正向前体乘波器工程设计方法进行了研究,给出了具体设计流程,并进行了初步的数值仿真验证,表明通过该方法设计的乘波前体流动特征与预期的结果吻合,说明文中所给出的激波附体条件及影响规律是可信的,乘波前体设计方法是可行的。     

斜激波串与上游激波干扰特性试验研究

为研究斜激波串在背压条件下前移与上游激波相互干扰的流场结构和运动规律,在来流为马赫数 2.7 的直管道内设计一种等宽度斜楔,采用动态压力测量、高速纹影和粒子图像测速(PIV)技术等手段进行了试验。研究结果表明：内置斜楔在管道内产生入射激波、分离激波、膨胀波、再附激波和激波诱导分离等复杂上游激波流场,在分离区附近形成有顺压梯度和逆压梯度的区域。当增大下游压比时,斜激波串逐渐向上游激波流场移动;经过斜楔产生的分离区时,斜激波串的移动速度急剧提升,同时出现非对称分离偏转方向的切换。对比了三种长度尺寸的等楔角斜楔所产生的上游激波流场的差异性,发现在相同的斜楔前缘起始点和楔角时,随着斜楔长度的增加,上游激波流场中激波诱导的分离尺度逐渐变大。     

一种进气道内激波/边界层干扰控制的新方法及其流动机理

针对高超声速进气道内经常存在的激波/边界层干扰现象,提出了一种基于可变形壁面鼓包的激波/边界层干扰控制概念,并对相关流动机理及参数影响规律进行了细致研究,结果表明：可变形鼓包通过其迎风侧的预增压作用,外凸段膨胀波束对反射激波的削弱作用,以及膨胀波束对边界层气流的加速作用来对激波/边界层干扰现象进行抑制;当激波入射点位于鼓包背风侧膨胀波区时,鼓包对边界层分离的抑制效果明显,并且适当增加鼓包高度可增加其抑制效果;对于鼓包迎风侧型线,在设计时应尽量采用较小的内凹段曲率,同时在外凸段上其最大曲率点应尽量与激波入射点靠拢,而对于背风侧型线的设计则应选择相近的外凸段和内凹段曲率较为合适。     

超燃冲压发动机隔离段内附面层/激波串相互干扰研究

隔离段内激波串的产生和发展以及激波/附面层相互干扰现象是极为复杂的,有效地进行激波串的组织是研究隔离段的关键所在,而其性能的好坏直接影响着超燃冲压发动机的性能。采用数值模拟的方法对不同来流附面层厚度工况的二维轴对称隔离段内流场流动特性进行了数值计算,分析了附面层/激波相互作用机理和附面层对隔离段激波串及隔离段性能的影响。结构表明:压缩-膨胀-再压缩-再膨胀……的气流流动挤压过程导致激波串的形成,逆压梯度的存在构成了附面层分离;附面层厚度的增加影响着激波串起始位置和结构;随着附面层厚度的增加,出口总压恢复系数和质量平均马赫数降低,且随着反压增大,变化趋势趋于明显。     

喉栓式推力可调发动机喷管流场数值模拟

对喉栓式推力可调固体火箭发动机喷管流场进行了数值模拟,并对喉栓型面进行了过程优化;针对喉栓不同作动速度和自由容积,分析了流场内各参数的变化;进行了非同轴喉栓发动机试验研究.计算结果表明,细长锥型喉栓总体性能最优;发动机压强建立过程与喉栓作动速度和自由容积关系密切;模拟结果与试验数据差别不大,可为喉栓式推力可调固体火箭发动机的研发提供参考.     

真实气体效应对高超声速轨道器气动特性的影响

基于一个7组元6反应动力学模型，采用NND差分格式求解化学反应Navier-Stokes方程，数值研究高超声速轨道器的绕流特性。重点讨论了轨道器气动特性在真实气体效应作用下对不同来流状态和不同舵偏角的敏感性。研究表明：真实气体效应主要发生在物面附近很薄的激波层内，缩短了激波的脱体距离，使激波层变薄，流动变量的梯度变大；空气的离解和电离导致轨道器的阻力系数比完全气体计算值低，压心位置前移。小攻角下，升力系数和俯仰力矩系数的真实气体计算值高于完全气体计算值，大攻角情形则相反。此外，小攻角时真实气体效应产生小低头力矩，而大攻角时产生小抬头力矩。单就舵面而言，真实气体效应使其阻力系数增大，使其升力系数和俯仰力矩系数在小攻角且非负舵偏角时变小，在大攻角且负舵偏角时变大。特别地，真实气体效应仅在零攻角且零舵偏角时对舵面的压心位置产生较大影响。     

高超声速乘波飞行器气动实验研究

以绕楔高超声速流场为基础，用流线追踪法生成了一种高超声速飞行器气动概念构形、初步探索了高超声速飞行器机身/推进系统一体化气动构形设计方法，开展了高超声速测压实验，结果表明：该类构形飞行器在高超声速飞行时，可以产生较高的升阻比，前体的预压缩效果明显，是以吸气式冲压发动机动力的有效途的飞行器构形。     

Growth and decay of weak discontinuities in radiative gasdynamics

Growth and decay properties of weak discontinuities headed by wave fronts of arbitrary shape in three dimensions are investigated in a thermally radiating inviscid gas flow. The effects of radiative transfer are treated by the use of a general differential approximation for a grey gas of arbitrary opacity including effects of radiative flux, pressure and energy density. The transport equations representing the rate of change of discontinuities in the normal derivatives of the flow variables are obtained, and it is found that the nonlinearity in the governing equations does not contribute anything to the radiation induced waves. In contrast to the radiation induced waves, the nonlinearity in the governing equations plays an important role in the interplay of damping and steepening tendencies of a modified gasdynamic wave. An explicit criterion for the growth and decay of a modified gasdynamic wave along bicharacteristics curve in the characteristic manifold of the governing differential equations is given and the special reference is made of diverging and converging waves. It is shown that there is a special case of a compressive converging wave for which the stabilizing influences of thermal radiation and the wave front curvature are not strong enough to overcome the tendency of the wave to grow into a shock.     

Asymmetric wedging of elastic material with the formation of separation zones

The problem of asymmetrical movement of a thin rigid impactor in elastic material is considered with subsonic and transonic velocities. For all considered range of velocities the flow diagram was determined for contours in the forms of wedge and ogive. There is a limiting speed value when the impactor moves with velocity higher than the speed of transverse wave at which the separation zone in the nasal part of the body disappears. At this velocity forces acting on the impactor are independent of its shape.     

Perturbation solutions for variable energy blast waves

The trajectory of and the flow field behind blast waves with time varying energy input is determined. Freeman's (1968) Lagrangean coordinate formulation is modified to include both the geometric factor, α, for plane, cylindrical and spherical shocks and also non-integer values of β, the energy input parameter, in a single computational algorithm. Numerical problems associated with vanishing density at the inner mass boundary or “piston face” are then examined and solved. Second order perturbation solutions about the solution for an infinite strength shock are then obtained in Sakurai's (1965) inverse shock Mach number expansion parameter for 0 β < α + 1. Tables and graphs of significant numerical coefficients are presented for comparison to, and extension of, results of other authors. Graphs of typical shock trajectories and flow field density, pressure and velocity variations are also presented and discussed.     

喉部曲率半径对最大推力喷管性能的影响

针对不同喉部结构进行了最大推力喷管型面设计计算,分析了不同曲率半径对喷管流量系数、几何效率、推力系数、扩张损失和分离点位置的影响。结果表明：流量系数随上游曲率半径的增大而增大,推力系数和几何效率随下游曲率半径的增大而减小;下游曲率半径的增大导致扩张损失在小范围内不断减小,分离点位置后移。     

前向空腔的弓形激波特性研究

在FD-14A激波风洞中Ma=10流场对前向空腔构型开展试验研究,应用高速阴影技术捕捉弓形激波的平均位置及振荡幅值,利用压力传感器测量空腔底部的脉动压力。在现有无空腔钝头体激波脱体距离预测方法的基础上发展了前向空腔构型的激波脱体距离预测方法,结合国外的试验测量结果与Organ-pipe理论,验证了这种方法的有效性和适用性,且该方法对激波脱体距离的预测结果与FD-14A风洞试验结果一致。此外,基于这种方法讨论了空腔振荡频率预测方法存在的争议。最后,研究了Ma=10流场下球锥体-前向空腔构型的脱体激波振荡幅值与平均速度的规律。     

激波在狭缝中绕射过程研究

针对固体火箭发动机药柱存在裂纹的情况，对运动激波在狭缝中绕射传播的非定常流场进行了计算，得到了清晰的流场变化图谱，从流场图中可分辨出激波在狭缝中的传播过程。对不同深宽比的狭缝进行研究，发现流场结构存在明显差别，在深宽比较小时狭缝内压力波振荡较小，当深宽比超过一定值时狭缝内的压力波产生剧烈振荡。研究结果表明，该方法可很好地描述激波在狭缝中的传播过程。     

可控曲率半径的圆柱形叶片逐点绘型方法

针对中低比转速离心泵,根据叶片进出口边界条件,以逐点绘型方法为基础,提出了一种新的曲率半径可控的叶片绘型方法。该方法的主要特点是曲率半径比值可作为设计常量由设计人员根据需要事先给定,随后分析了曲率半径及比例因子对叶片安放角、叶片包角、相对速度及速度矩等的影响。结果表明,不同曲率半径比值下的叶型参数及流动参数变化范围很大,曲率半径比值较大时,节流损失较大,泵扬程较低,曲率半径比值较小时,脱流损失较大,泵效率较低,存在较优的曲率半径比值区间[1.4,2.4],使叶片安放角平滑变化,泵的综合性能较优,在该优化区间内,取较大的曲率半径比值有利于获得较优的汽蚀性能,比例因子为0时叶片安放角的变化较为平稳,可用于开展离心泵的初步设计。     

Transient analysis of counterflowing jet over highly blunt cone in hypersonic flow

Understanding the characteristics of various Counterflowing jets exiting from a nose cone is crucial for determining heat load reduction and usage of this device in various conditions. Such jets can undergo several flow regimes during venting, from initial supersonic flow, to transonic, to subsonic flow regimes as the pressure of jet decreases. A bow shock wave is a characteristic flow structure during the initial stage of the jet development, and this paper focuses on the development of the bow shock wave and the jet structure behind it. The transient behavior of a sonic counterflow jet is investigated using unsteady, axisymmetric Navier–Stokes solved with SST turbulence model at free stream Mach number of 5.75. The coolant gas (Carbon Dioxide and Helium) is chosen to inject into the hypersonic air flow at the nose of the model. The gases are considered to be ideal, and the computational domain is axisymmetric. The jet structure, including the shock wave and flow separation due to an adverse pressure gradient at the nose is investigated with a focus on the differences between high diffusivity coolant jet (Helium) and low diffusivity coolant jet (CO₂) flow scenarios.     

“人字形小肋”对激波/边界层干扰的影响研究

在曼彻斯特大学跨声速风洞开展激波/边界层干扰及“人字形小肋”对其影响的实验研究。在马赫数1.85流场条件下,应用高速纹影、油流、皮托压力测量和基于压敏漆的壁面压力测量技术,研究“人字形小肋”流动控制方法对激波/边界层干扰的流动分离结构与尺寸、压力分布特性与波系特征等影响。结果显示激波/边界层干扰诱发流动分离,分离区呈现三维特征,在“人字形小肋”的作用下,分离线呈现“波浪”形且整体向上游移动,干扰区流向尺寸增大,分离区高度减小且长度略增大,再附区的压力极值降低,这些特征与叶片、尖楔等微涡发生器的影响趋势相反。下一步工作中,拟针对“人字形小肋”开展参数优化研究,“人字形小肋”可能成为降低激波/边界层干扰诱发的高热流载荷的有效方法。     

A numerical study of the charge diameter effect on unsteady detonation

A numerical study for the unsteady detonation of an unconfined tetryl charge of small diameter, which is assumed to be homogeneous, was performed by using the two-dimensional Lagrangian hydrodynamic computer code, 2 DL, with the first order Arrhenius equation of reaction rate. Becker-Kistiakowsky-Wilson (BKW) and Kihara-Hikita (KH) equations of state have been applied to the detonation products.In the case of BKW, it is shown that the rarefaction waves propagating inward from the lateral surface make the reaction rate slow and give a curvature to the front. Then after an induction time, a strong initiation occurs in the reaction zone near the lateral surface and higher pressure zone moves to the axis. This higher pressure accelerates the detonation propagation near the lateral surface and the curvature of detonation front is reduced. Then, the reaction at the lateral surface again begins to decay by the rarefaction waves. Such a sequence of process is repeated periodically.The possibility of the occurrence of the strong initiation depends on the pressure and temperature in the shocked zone near the surface. In a small diameter charge, the delayed explosion becomes weaker near the surface, while its frequency increases. No shock interaction occurs because the direction of the particle flow is always divergent.In the case of KH equation of state, the temperature of detonation is higher than that obtained by BKW and the behaviour of instability becomes rather different from the previous result, i.e. in the axis the pressure oscillates repeating the overdriven and underdriven detonation similar with the case of BKW.     

Numerical validation and parametric investigation on the cold flow field of a typical cavity-based scramjet combustor

The three-dimensional coupled implicit Reynolds Averaged Navier–Stokes (RANS) equations and the two equation standard k–ε turbulence model has been employed to numerically simulate the cold flow field in a typical cavity-based scramjet combustor. The numerical results show reasonable agreement with the schlieren photograph and the pressure distribution available in the open literature. The pressure distribution after the first pressure rise is under-predicted. There are five shock waves existing in the cold flow field of the referenced combustor. The first and second pressure rises on the upper wall of the combustor are predicted accurately with the medium grid. The other three shock waves occur in the core flow of the combustor. The location of the pressure rise due to these three shock waves could not be predicted accurately due to the presence of recirculation zone downstream of the small step. Further, the effect of length-to-depth ratio of the cavity and the back pressure on the wave structure in the combustor has been investigated. The obtained results show that there is an optimal length-to-depth ratio for the cavity to restrict the movement of the shock wave train in the flow field of the scramjet combustor. The low velocity region in the cavity affects the downstream flow field for low back pressure. The intensity of the shock wave generated at the exit of the isolator depends on the back pressure at the exit of the combustor and this in turn affects the pressure distribution on the upper wall of the combustor.