Experimental Investigation of a Fixed-geometry Two-dimensional Mixed-compression Supersonic Inlet with Sweep-forward High- light and Bleed Slot in an Inverted ＂X＂-type Layout

A fixed-geometry two-dimensional mixed-compression supersonic inlet with sweep-forward high-light and bleed slot in an inverted ＂X＂-form layout was tested in a wind tunnel. Results indicate： （1） with increases of the free stream Mach number, the total pressure recovery decreases, while the mass flow ratio increases to the maximum at the design point and then decreases; （2） when the angle of attack, a, is less than 6°, the total pressure recovery of both side inlets tends to decrease, but, on the lee side inlet, its values are higher than those on the windward side inlet, and the mass flow ratio on lee side inlet increases first and then falls, while on the windward side it keeps declining slowly with the sum of mass flow on both sides remaining almost constant; （3） with the attack angle, a, rising from 6° to 9°, both total pressure recovery and mass flow ratio on the lee side inlet fall quickly, but on the windward side inlet can be observed decreases in the total pressure recovery and increases in the mass flow ratio; （4） by comparing the velocity and back pressure characterristics of the inlet with a bleed slot to those of the inlet without, it stands to reason that the existence of a bleed slot has not only widened the steady working range of inlet, but also made an enormous improvement in its performance at high Mach numbers. Besides, this paper also presents an example to show how this type of inlet is designed.

CHAOTIC MOTIONS AND LIMIT CYCLE FLUTTER OF TWO-DIMENSIONAL WING IN SUPERSONIC FLOW

Based on the piston theory of supersonic flow and the energy method, the flutter motion equations of a two-dimensional wing with cubic stiffness in the pitching direction are established. The aeroelastic system contains both structural and aerodynamic nonlinearities. Hopf bifurcation theory is used to analyze the flutter speed of the system. The effects of system parameters on the flutter speed are studied. The 4th order Runge-Kutta method is used to calculate the stable limit cycle responses and chaotic motions of the aeroelastic system. Results show that the number and the stability of equilibrium points of the system vary with the increase of flow speed. Besides the simple limit cycle response of period 1, there are also period-doubling responses and chaotic motions in the flutter system. The route leading to chaos in the aeroelastic model used here is the period-doubling bifurcation. The chaotic motions in the system occur only when the flow speed is higher than the linear divergent speed and the initial condition is very small. Moreover, the flow speed regions in which the system behaves chaos axe very narrow.

Flow response hysteresis of throat regulation process of a two-dimensional mixed-compression supersonic inlet

The variable geometry supersonic inlet tends to decrease the throat area to reduce the Mach number upstream of the terminal shock,so as to reduce the flow loss.However,excessive Internal Contraction Ratio(ICR)exposes the inlet to a greater risk of unstart,which inevitably results in a process of increasing the throat area to aid the inlet restart.In the above throat regulation process,the inlet undergoes the start,unstart,and restart states in turn.In order to reveal the flow structure and mechanism of this process,a two-dimensional unsteady numerical simulation combined with a dynamic mesh technique were employed.The shock-on-lip Mach number of the studied inlet is 4.0 and the flight angle of attack is+6°.Analysis was focused on the state with a freestream Mach number of 3.0.The results clearly show that the flow response hysteresis appears,and restart is only realized when the throat area is obviously increased as compared to that of unstart due to the historical unstart flow structure.In addition,three typical flow fields were analyzed,and it is found that the separation ahead of the inlet was the key factor affecting the hysteresis.Finally,unstart and restart boundaries of the inlet were discussed,and the factors influencing its deviation from the typical boundaries of dual-solution area were analyzed.The newly predicted unstart and restart boundaries are much closer to the CFD results.

超声速进气道出口弯段的阻力特性数值研究

推阻匹配设计是吸气式飞行器设计的核心问题,为了获取进气道的阻力特性并降低内流阻力,提高吸气式飞行器的总体性能,采用数值方法对超声速双侧布局进气道的冷流阻力特性开展研究,对比分析了不同转弯长度、转弯角度、扩张比下的弯段流态和阻力特性;获得了进气道内阻的分配比例及关键几何参数对弯段流场结构和进气道内阻的影响特性。结果表明:对于双侧布局的进气道,冷流条件下弯段内容易发生流动分离,存在较大的流动损失,在不产生溢流的情况下,弯段内阻占整个进气道内阻的大部分。分析发现,减小转弯角度或增加弯段扩张比均可降低进气道弯段内阻,而转弯长度与转弯半径相互影响,在给定的设计条件下,弯段内阻随转弯长度的增加先减小后增大。

二元超声速混压式进气道亚临界稳定裕度研究

为了研究内收缩比和来流马赫数对二元超声速混压式进气道亚临界稳定裕度的影响规律及失稳机制,采用二维非定常仿真方法研究了内收缩比(ICR)为1.04~1.25的进气道在来流马赫数Ma0为2.4的条件下,以及内收缩比为1.08的进气道在来流马赫数为2.2~2.8条件下,其由稳态向失稳状态转变的过程。研究结果表明:(1)当Ma0=2.4时,在1.04≤ICR≤1.12内,随着ICR增加,亚临界稳定裕度ζ减小;1.16≤ICR≤1.25内,随着ICR增加,亚临界稳定裕度增大。(2)在内收缩比为1.08的条件下,马赫数变化引起的分离激波角和分离包再附压升两个关键因素变化共同主宰着进气道亚临界稳定裕度的变化趋势。(3)总体上,根据稳定亚临界初始状态的三相点无量纲高度?b是否大于1可将进气道的亚临界稳定裕度变化情形分为两类,当?b<1时,ζ随着?b的增加而减小;当?b> 1时,ζ随着?b的增加而增加。

Aerodynamic Design Method and Experimental Investigation of a High-Load Supersonic Compressor Cascade

The high-load compressor plays an important role in further improving the performance of aero-engine.However,the complex shock waves in the cascade channel also bring more aerodynamic losses.This paper proposes a supersonic compressor cascade modeling method based on the theory of unique inlet flow angle,and the aerodynamic design and optimization of a cascade with inlet Mach number 1.85 are studied by combining the numerical optimization method and planar cascade experiment.The results show that pressure increase can be achieved by multiple shock waves which are obtained by the reflection of the leading edge detached shock wave in the initial supersonic cascade channel at the design point,which verifies the feasibility of the design method.After optimization,the aerodynamic performance of the cascade has been improved to different degrees at the design point and off-design point.When the static pressure ratio is 3.285,the total pressure recovery coefficient reaches 86.82%at the design point,which is on the advanced level of the same type of cascade.The experimental results of planar cascade schlieren and surface pressure measurement also verify the correctness of the simulation method,which provides useful references for the subsequent compressor design.

Experimental investigation on unstart-restart hysteresis of a supersonic inlet during throat regulation

The hysteresis during the throat regulation process of a supersonic variable inlet is unconducive to restart.Hence,detailed experimental studies of such a hysteresis and its control are necessary.A throat variable supersonic inlet was designed at a shock-on-lip Mach number of 4.0 and an Internal Contraction Ratio(ICR)ranging over 1.21–2.94.Meanwhile,a distributed bleed system was proposed to suppress the hysteresis.The wind tunnel tests were conducted at Mach number 2.9.The throat regulation processes were recorded using a high-speed schlieren and dynamic pressure acquisition system.The results indicate that the unstart and restart ICRs during the uncontrolled inlet’s throat regulation process were 1.95 and 1.48,respectively,demonstrating an unstart-restart hysteresis.Four typical flowfields were summarized during the uncontrolled inlet’s restart process.The proposed bleed control increased the unstart and restart ICRs to 2.06 and 1.75,respectively,and the inlet realized the designed state as the ICR was further decreased to 1.67.The controlled inlet’s hysteresis loop was decreased compared to the uncontrolled inlet.Finally,the mechanism of the hysteresis,dominated by the entrance separation-induced wave system,was clarified.The mechanisms of the bleed control to broaden the unstart and restart boundaries and suppress the hysteresis were elucidated.

The Two-Dimensional Supersonic Flow and Mixing with a Perpendicular Injection in a Scramjet Combustor

A numerical investigation has been performed on supersonic mixing of hydrogen with air in a Scramjet (Supersonic Combustion Ramjet) combustor and its flame holding capability by solving Two-Dimensional full Navier-Stokes equations. The main flow is air entering through a finite width of inlet and gaseous hydrogen is injected perpendicularly from the side wall. An explicit Harten-Yee Non-MUSCL Modified-flux-type TVD scheme has been used to solve the system of equations, and a zero-equation algebraic turbulence model to calculate the eddy viscosity coefficient. In this study the enhancement of mixing and good flame holding capability of a supersonic combustor have been investigated by varying the distance of injector position from left boundary keeping constant the backward-facing step height and other calculation parameters. The results show that the configuration for small distance of injector position has high mixing efficiency but the upstream recirculation can not evolved properly which is an important factor for flame holding capability. On the other hand, the configuration for very long distance has lower mixing efficiency due to lower gradient of hydrogen mass concentration on the top of injector caused by the expansion of side jet in both upstream and downstream of injector. For moderate distance of injector position, large and elongated upstream recirculation can evolve which might be activated as a good flame holder.

超声速进气道与脉冲爆震燃烧室匹配工作特性研究

为探索设计点马赫数为2.5的三波系二元混压式超声速进气道与脉冲爆震燃烧室的匹配工作特性,设计了3种工作方案,采用带有化学反应的非定常数值仿真方法,研究了多循环工作过程。结果表明:燃烧室出口泄压与燃烧室进口机械阀均能抑制反传压缩波向上游的传播,拓宽进气道稳定工作裕度,且能提高循环频率;但前者会损失推力,而后者能大幅增加推力。机械阀会对燃气中的压缩波产生反射作用,增大燃烧室内湍流度,使得在开阀瞬间燃气被卷吸至上游扩张段内,但燃气中的反射压缩波不会影响爆震波波速与形态。在爆震波起爆与传播过程中,燃烧室进口应尽可能扩大物理遮挡物阻塞比,且不应对进气道出口流动造成堵塞,增加流动阻力。

超声速进气道唇口剖面形状对飞机升阻比的影响

针对一种采用无隔道超声速进气道的超声速靶机,使用数值仿真方法对内外流一体化的全机流场进行数值模拟,研究不同唇口剖面形状对全机升阻比的影响。结果表明:设计状态下不同唇口剖面形状升阻比不同。综合考虑高低速飞行马赫数在0°~6°飞行攻角下,飞机的气动力特性,长短轴比为9∶1的椭圆唇口为最佳唇口剖面形状。

Development of a coupled supersonic inlet-fan Navier–Stokes simulation method

A coupled supersonic inlet-fan Navier–Stokes simulation method was developed by using COMSOL-CFD code. The flow turning, pressure rise and loss effects across blade rows of the fan and the inlet-fan interactions were taken into account as source terms of the governing equations without a blade geometry by a body force model. In this model, viscous effects in blade passages can also be calculated directly, which include the exchange of momentum between fluids and detailed viscous flow close to walls. NASA Rotor 37 compressor test rig was used to validate the ability of the body force model to estimate the real performance of blade rows. Calculated pressure ratio characteristics and the distribution of the total pressure, total temperature, and swirl angle in the span direction agreed well with experimental and numerical data. It is shown that the body force model is a promising approach for predicting the flow field of the turbomachinery. Then, coupled axisymmetric mixed compression supersonic inlet-fan simulations were conducted at Mach number 2.8 operating conditions. The analysis includes coupled steady-state performance, and effects of the fan on the inlet. The results indicate that the coupled simulation method is capable of simulating behavior of the supersonic inlet-fan system.

Axisymmetric and three-dimensional flow simulation of a mixed compression supersonic air inlet

The flow through an axisymmetric supersonic mixed-compression air inlet has been simulated numerically to investigate the effects and the necessity of the three-dimensional(3D)modeling in comparison with the axisymmetric one.For this purpose,a supersonic inlet has been simulated numerically via axisymmetric and 3D CFD solvers,using the steady state Reynolds-averaged Navier-Stokes equations along with the SST k-ωturbulence model,for a free-stream Mach number of 2.0 and at zero degrees angle of attack.The grid for the 3D cases was a 14.4-degree sector,instead of a 360-degree domain one,with rotational periodic boundary condition for the side boundaries.The results show that both static and total pressure distributions match well with the experimental data for both the axisymmetric and the 3D simulations.If the prediction of performance parameters is the main goal of simulations,it seems that the axisymmetric simulation provides adequate accuracy,and the 3D simulation one is not the best choice.The 3D numerical simulation results in an in-depth study on the supersonic inlets,including the shock wave-boundary layer interaction,the location of the terminal normal shock wave,and consequently the separation point.For an axisymmetric supersonic inlet in an axisymmetric flow condition,3D effects are not strong enough to have a significant influence on the inlet performance for all operational conditions.However,it seems that 3D effects play an important role in both critical and supercritical operating conditions during the steady state operation.

Minimization of Classification Samples for Supercritical and Subcritical Patterns of Supersonic Inlet

In order to investigate sample minimization for classification of supercritical and subcritical patterns in supersonic inlet, three optimization methods, namely, opposite one towards nearest method, closest one towards the byper-plane method and random selection method, are proposed for investigation on minimization of classification samples for supercritical and subcritical patterns of supersonic inlet. The study has been carried out to analyze wind tunnel test data and to compare the classification accuracy based on those three methods with or without priori knowledge. Those three methods are different from each other by different selecting methods for samples. The results show that one of the optimization methods needs the minimization samples to get the highest classification accuracy without priori knowledge. Meanwhile, the number of minimization samples needed to get highest classification accuracy can be further reduced by introducing priori knowledge. Furthermore, it demonstrates that the best optimization method has been found by comparing all cases studied with or without introducing priori knowledge. This method can be applied to reduce the number of wind tunnel tests to obtain the inlet performance and to identify the supercritical/subcritical modes for supersonic inlet.

超声速进气道喉部附面层抽吸

为研究超声速进气道喉部之后流场激波附面层干扰,采用FLUENT软件模拟了单楔角进气道在设计工况下流动情况。通过分析,提出进气道喉部抽吸。计算了三种抽吸缝大小下进气道喉部之后流场,计算结果表明,喉部抽吸能使激波稳定于喉部,通过抽吸能改善喉部之后流场状况,提高进气道性能,少量抽气不改变流场结构,加大抽气量,使喉部之后激波串转变成正激波,正激波之后流场不分离,进气道出口性能参数提高显著。

侧压式进气道与飞行器机体气动一体化设计及实验

以机体 推进系统耦合、三维侧压式进气道为基本特征,设计了采用超燃冲压发动机为推进系统的高超声速一体化冷流通气实验模型,在高超声速炮风洞中完成了飞行器的整体气动测力试验。在来流马赫数Ma=8.09的条件下,分别测定了飞行器结合单模块、3模块、5模块超燃冲压发动机在-4°～6°六个攻角下模型的气动力数据,并对实验结果作了分析。

凸包(Bump)进气道/DSI模型设计及气动特性研究

利用基于乘波理论的Bum p进气道设计软件,据锥型流精确流线法设计了一腹部进气布局的Bum p进气道。用CFD模拟手段,从对称面马赫数分布、凸包上的极限流线及横截面上压强系数分布等方面分析了进气道进口段流场特征,证明所设计的Bum p进气道流动特征符合预期设计目标。通过CFD计算和试验对比,分析了所设计的Bum p进气道超声速气动性能,表明在发动机设计状态,在来流马赫数M a∞为2.0时,出口平均总压恢复系数接近0.87,而在M a∞=1.8时该值不低于0.91。

碳氢燃料超燃冲压发动机进气道与燃烧室匹配性能试验研究

在模拟飞行马赫数Ma=6,高度25km条件的液体碳氢燃料超燃冲压发动机自由射流试验中,对比研究了4种不同进气道,不同燃烧室入口条件下模型发动机的点火与燃烧性能。试验结果表明几何内收缩比3的侧压式进气道的出口压强低而无法实现模型发动机的点火;进气道增加部分前体压缩,模型发动机则能够维持稳定燃烧,得到正推力;采用较高收缩比5 35的三维进气道的出口流场畸变程度较高,降低了隔离段抗反压的能力,会对燃烧性能产生很大影响,燃烧效率、发动机推力显著下降,甚至可能导致发动机熄火。不同长度的隔离段对比研究表明隔离段加长能够提高抗反压能力,有助于实现煤油分级燃烧,提高燃烧效率。

冲压发动机进气道压力振荡过程的数值研究

针对来流马赫数3.5,0°迎角飞行条件下冲压发动机的简化模型,数值模拟了进气道整流罩开启过程的非定常流动。计算采用AUSMPW有限体积格式,基于结构与非结构的混合网格技术,求解了非定常N-S方程。计算结果表明,在进气道整流罩打开以后,超声速气流会在燃烧室通道内形成振幅衰减较慢的低频剧烈振荡。在同样计算条件下采用Euler方程进行对比性研究,数值实验发现对这类大雷诺数、强非定常性的问题,Euler方程和N-S方程得到的结果相差不大。

宽Ma数范围固定几何进气道设计问题研究

针对Ma=2～3.5中等超声速宽范围的设计要求,通过实例方案设计和CFD数值模拟分析表明,按工程上现行的折中设计思想,固定几何进气道的内流道型面实质上是为满足低速接力要求而设计的。在高速巡航时,其超声速扩压段气流总折转角偏小,且喉道高度偏大,对捕获来流压缩不充分,导致喉道段的气流流速过高,且畸变显著,致使结尾正激波总压损失过大,是进气道高速性能差的主要成因。

旋转冲压压缩转子二维进气流道数值研究

借鉴二维超声速进气道的设计方法,设计了一种旋转冲压压缩转子的二维进气流道,并采用二维雷诺平均N-S方程和Spalart-Allmaras湍流模型对其流场进行了数值仿真,研究了转速、背压对二维进气流道中波系结构、内部流动特性和性能的影响.计算结果表明:所提出的旋转冲压压缩转子二维进气流道设计方法是可行的;二维进气流道中产生的激波系与二维超声速进气道中产生的激波系相类似,所不同的是二维进气流道中产生的激波为弯曲激波,而二维超声速进气道中产生的激波为平直激波;转速和背压对二维进气流道的性能有较大影响.