Application of Nano Technique in Measuring Supersonic/Hypersonic Flow

Turbulence, universally exist in nature and human activities, is a kind of three-dimensional, irregular, unsteady flow. Ever since 19th century when people started to investigated turbulent flow technically, they have never dropped the po-tent and intuitionistic experimental method. Recently, with the development of aviation and aerospace industry, espe-cially with the increase desire of supersonic and hypersonic flight, the mechanism of high speed and compressible flow has become hot topic of fluid research, resulting in development of measurement method and technique. When encoun-tering compressible high flow, traditional techniques, such as schilieren, shadow and interference, cannot measure fine flow structures. Fortunately, multiple-discipline integration of nano technique, laser technique and imaging technique provides a new design for fluid measurement。Nano-tracer planar laser scattering (NPLS) is a new flow visualization technique, which was developed by the authors’ group in 2005, and it can visualize time correctional flow structure in a cross-section of instantaneous 3D supersonic flow at high spatiotemporal resolution. Many studies have demonstrated that NPLS is a powerful tool to study supersonic turbulence.

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.

Robust adaptive control of hypersonic vehicle considering inlet unstart

In this paper,a model reference adaptive control(MRAC)augmentation method of a linear controller is proposed for air-breathing hypersonic vehicle(AHV)during inlet unstart.With the development of hypersonic flight technology,hypersonic vehicles have been gradually moving to the stage of weaponization.During the maneuvers,changes of attitude,Mach number and the back pressure can cause the inlet unstart phenomenon of scramjet.Inlet unstart causes significant changes in the aerodynamics of AHV,which may lead to deterioration of the tracking performance or instability of the control system.Therefore,we firstly establish the model of hypersonic vehicle considering inlet unstart,in which the changes of aerodynamics caused by inlet unstart is described as nonlinear uncertainty.Then,an MRAC augmentation method of a linear controller is proposed and the radial basis function(RBF)neural network is used to schedule the adaptive parameters of MRAC.Furthermore,the Lyapunov function is constructed to prove the stability of the proposed method.Finally,numerical simulations show that compared with the linear control method,the proposed method can stabilize the attitude of the hypersonic vehicle more quickly after the inlet unstart,which provides favorable conditions for inlet restart,thus verifying the effectiveness of the augmentation method proposed in the paper.

Analysis of Unstarted Hypersonic Flow Unsteadiness Based on Schlieren Image Processing

The unstarted flow field in a hypersonic inlet model at a design point of Ma 6 is studied experimentally.The time-resolved spatial flow characteristics of the separation shock oscillation,which is induced by the unstarted flow,are analyzed based on a high-speed Schlieren system and an image processing method.The motion of the separation shock detected by the shock-detection algorithm is compared to the results of fast-response wall-pressure measurements,and good agreement is demonstrated by comparing the frequency components in the power spectral density contours between shock oscillation and pressure fluctuation.The hysteresis of the pressure and separation shock during oscillation cycles is observed from the time history of the shock motion,which means that the unsteady flow pattern of the unstarted hypersonic flow can be accurately clarified by time-resolved Schlieren image processing.These results convincingly demonstrate that the shock-detection technique is successfully applied to an unstarted hypersonic flow case.

FOREBODY COMPRESSIBILITY RESEARCH OF HYPERSONIC VEHICLE

Three kinds of forebody model of hypersonic vehicles were studied with numerical simulation method. It shows that the two-order compressive ramp model is the best selection among the three for its good evaluative parameters value at the cowl of the inlet. This model can provide higher value of flux coefficient and total pressure recovery coefficient and lower average Mach number compared with those of the other two models. Simultaneously different compressive angles may have different effects. The configuration which the first-order of compressive angle is 4° and the second 5° is the optimum combination. Furthermore factors such as attack angle were concerned. Better result may be obtained with a range of attack angles. Based on the work above the integrated design for forebody/inlet of a hypersonic vehicle was performed. The numerical result shows that this integrated model provides good flow field quality for inlet and engine work.

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。

氢预冷涡轮发动机研究进展及关键技术

高速涡轮发动机及其组合动力装置是高超声速飞行器技术的基础和关键。随着飞行马赫数提高,来流空气总温显著升高,发动机推力急剧减小。在此背景下,进气预冷成为扩展航空涡轮发动机工作速域的主要方向。液氢同时具备高热值和高热沉,是燃料换热预冷的理想工质。因此,氢预冷涡轮发动机被视为实现临近空间高超声速飞行的重要技术之一。回顾了国外氢预冷吸气式发动机的发展历程,分析了各型发动机的主要特点,并根据预冷目的归纳总结了面向氢氧火箭以及面向冲压或涡喷发动机的2类氢预冷技术。在此基础上,考虑氢预冷涡轮发动机的工作需求,对其研发中的关键技术进行了梳理。与传统航空发动机相比,氢预冷涡轮发动机由于采用了新的循环、燃料和结构,给总体、传热、燃烧、材料等方面带来了诸多挑战。其中的关键技术包括:预冷系统与发动机总体性能的全工况稳态和动态匹配技术;高功重比预冷器的设计、成型和防冰技术;氢燃料动态高精度计量和燃烧控制技术;涉高压氢部件的氢损伤抑制及预测技术等。

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。

燃料预喷注与释热对高马赫数进气道性能影响的数值研究

为掌握高马赫数条件下燃料预喷注与释热对进气道性能的影响,以来流马赫数Ma_(∞)=10为设计点进行了典型二元构型进气道设计与基准流场仿真,开展了Ma_(∞)=7~10进气道氢燃料预喷注、反应流场仿真研究。结果表明:预喷注引起局部压缩激波系变化、改变下游流动特性、从而影响进气道性能,主要会导致总压恢复系数、喉部马赫数减小;同时预喷注动作具有一定的激波系调节、压缩循环调控的作用,且与预喷注位置密切相关,外压缩喷注会诱导外部斜激波向亚额定偏移,导致流量溢流增加、流量系数减小;内压缩喷注可以避免流量溢流,对进气道性能影响相对较小,且双侧组合喷注的预混效率可达0.60以上。反应流场仿真显示,燃烧反应主要发生在近壁面区域,流向上主要释热区在内压缩段,反应释热进一步降低总压恢复系数、减小喉部马赫数,外压缩喷注时会增加流量溢流,而循环静压比、静温比均相对增大,阻力系数基本保持不变,整体上预喷注及释热影响效应具有可控性。

基于膨胀波效应的高超声速进气道肩部流动分离控制研究

为改善高超声速进气道唇口激波/附面层干扰诱导的肩部流动分离,从膨胀波及激波理论出发,推导出了膨胀波效应影响下的斜激波附面层干扰理论公式,获得了影响斜激波诱导分离的主要因素:膨胀角梯度、激波角及波前马赫数。在此基础上,开展了膨胀波效应影响下的流动分离控制研究,给出了膨胀波效应影响下斜激波诱导分离的判别及预测方法。结果表明:增大激波入射点处膨胀角梯度,可以显著减小甚至消除肩部流动分离;随着激波角增大,激波强度及逆压力梯度增加,分离区尺寸显著增大。而波前马赫数对分离区尺寸的影响不显著;在进口马赫数3.57~5.18,唇罩角度6°~10°范围内,当激波入射点处逆压比梯度小于250 m^(-1)时,斜激波诱导的流动分离消失,可为改善超声速/高超声速进气道内流道流动分离提供技术支撑。

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.

Active control of supersonic/hypersonic aeroelastic flutter for a two-dimensional airfoil with flap

The flutter, post-flutter and active control of a two-dimensional airfoil with control surface operating in supersonic/hypersonic flight speed regions are investigated in this paper. A three-degree-of-freedom dynamic model is established, in which both the cubic nonlinear structural stiffness and the nonlinear aerodynamic load are accounted for. The third order Piston Theory is employed to derive the aerodynamic loads in the supersonic/hypersonic airflow. Nonlinear flutter happens with a phenomenon of limit cycle oscillations (LCOs) when the flight speed is less than or greater than linear critical speed. The LQR approach is employed to design a control law to increase both the linear and nonlinear critical speeds of aerodynamic flutter, and then a combined control law is proposed in order to reduce the amplitude of LCOs by adding a cubic nonlinear feedback control. The dynamic responses of the controlled system are given and used to compare with those of the uncontrolled system. Results of simulation show that the active flutter control method proposed here is effective.

Investigation of the effect of nozzle shape on supersonic/hypersonic impactors designed for size discrimination of nanoparticles

In this study the flow field and the nanoparticle collection efficiency of supersonic/hypersonic impactors with different nozzle shapes were studied using a computational modeling approach. The aim of this study was to develop a nozzle design for supersonic]hypersonic impactors with the smallest possible cut-off size d5o and rather sharp collection efficiency curves. The simulation results show that the changes in the angle and width of a converging nozzle do not alter the cut-off size of the impactor; however, using a conical Laval nozzle with an L]Dn ratio less than or equal to 2 reduced d5o. The effect of using a cap as a focuser in the nozzle of a supersonic/hypersonic impactor was also investigated. The results show that adding a cap in front of the nozzle had a noticeable effect on decreasing the cut-off size of the impactor. Both fiat disks and conical caps were examined, and it was observed that the nozzle with the conical cap had a lower cut-off size.

Computational investigation of powder coating of nanoparticles in supersonic and hypersonic impactors

In this study, numerical simulation of flow field in a supersonic/hypersonic impactor with one or two nozzles was carried out using a commercial computational fluid dynamics （CFD） software FLUENT. The objective was to investigate the effects of working parameters such as pressure ratio （50 〈 Po]Pb 〈 800）, nozzle diameters （D=0.23, 0.27, 0.45 mm）, nozzle to plate distance （0.5 〈L/D〈 50）, particle diameter （1 nm〈 dp 〈 100 nm ） and angle between two nozzles. A single-phase 3D unsteady-state model was implemented by the software. For this purpose, a user-defined function （UDF） was employed to implement nanoparticles for different assumptions of Cunningham correction factor. An axisymmetric form of the compressible Navier-Stokes and energy equations was used for both fluid flow and temperature; Lagrangian particle trajectory analysis was used for particle motion. Using the variable Cunningham cor- rection factor showed suitable agreement with experimental data in comparison with other methods. Results show that increase of the distance between nozzle and impaction plate causes increase of Mach number, the distance between bow shock and impaction plate, and the collection efficiency. Maximum jet velocity, distance between bow shock and impaction plate and collection efficiency increase by using two nozzles in supersonic and hypersonic imoactors.

高超声速进气道复杂内流热气动弹性研究

高超声速进气道在复杂波系的气动载荷和气动热作用下非常容易诱发热气动弹性问题,深入理解复杂内流下热气动弹性机理对未来高超声速进气道的精细化设计具有重要意义。建立了静/动热气动弹性动力学分析框架,深入研究了静/动热气动弹性对三维高超声速进气道流场结构和性能影响的规律和机理。静热气动弹性分析结果表明,双向耦合方法得到的气动热弹性变形相对较大,入口唇前缘变形量最大。结构变形改变了唇缘附近的激波结构,增强了进气道内部的激波强度,增加了分离区长度和外壁面温度,改变了出口流场。同时,热气动弹性变形会导致质量流量系数和压升比的增大,降低了总压恢复系数。动热气动弹性分析结果表明,对于模型,不考虑气动加热时,结构位移响应逐渐呈现收敛趋势;考虑气动加热后,结构位移响应呈现极限环的趋势。气动加热可能会改变进气道结构动态响应特征。由于进气道结构频率非常接近,结构动力响应中存在着“拍”现象。前缘变形较大而振幅较小,尾缘变形较小而振幅较大。结构振动导致流场结构产生明显的动态变化,且导致性能参数存在明显的波动,尤其是出口反压比波动幅度较大。希望通过研究加深对进气道中复杂波系结构中热气动弹性问题的理解与认识,以期为未来进气道的精细化设计提供参考。

预冷器对高超声速轴对称进气道设计状态气动性能影响

为获取预冷器对高超声速进气道内流特性的影响机理和影响规律,设计了一种在扩张段加入台阶型预冷器的高超声速轴对称进气道,并利用混合网格建立了仿真模型,获得了Ma6.0来流条件下带预冷器进气道与原型进气道在节流状态的数值仿真结果。结果表明:引入预冷器后进气道临界耐反压能力略有下降,并且进气道下游背压在1.0≤p_(b)/p_(0)<150时出口性能参数明显下降;预冷器上游总是存在节流,上游节流程度由预冷器的堵塞和出口背压共同决定,当p_(b)/p_(0)≤20时,上游流场结构完全由预冷器的堵塞作用决定,当p_(b)/p_(0)>20时,由两者共同决定;进气道下游背压在20≤p_(b)/p_(0)<275时,预冷器为上游流场带来消极影响,而当背压在20≤p_(b)/p_(0)<200范围时,预冷器的引入能有效改善下游流场品质,通过上下游的耦合作用,出口性能参数在p_(b)/p_(0)≥150后与原型进气道趋于一致。当p_(b)/p_(0)≥275时,唇罩侧放气缝对激波串根部低能流的抽吸使得预冷器几乎不对上游产生影响。

基于离散等收缩比的前体/进气道流向双乘波一体化设计

前体/进气道一体化设计是高超声速飞行的关键技术,一体化设计的核心是前体与进气道在基准流场上的气动融合.针对腹部进气布局中前体压缩后的非均匀流影响进气道性能的问题,文章基于局部收缩比处处一致的思想,提出了离散等收缩比设计方法,实现了乘波前体/内转式进气道流向气动融合与遵循气动规律的变截面流道设计.将进气道的三维流场分解成一簇具有相同收缩比的三维流管,视每根流管侧壁为轴对称流场;以锥导乘波前体压缩后的非均匀流作为来流条件,以总压恢复为目标对每根流管进行优化设计;通过匹配激波反射位置将流管重新组合起来,流管的对应边界组成内转式变截面进气道.该设计方法适配任何已知的非均匀来流,可灵活控制唇口位置,且适用于任意形状之间的变截面转换.数值研究表明,依托该方法设计的一体化构型性能符合预期,出口流场均匀,具有优越的抗反压能力,且非设计点流场波系结构良好.离散等收缩比设计方法为腹部进气布局中前体/进气道一体化气动融合设计提供了新思路.

基于丝线流动显示技术的内转进气道起动性能实验

采用丝线法流动显示技术,在高超声速冷流暂冲式下吹风洞开展了快速获取内转进气道起动性能的实验研究。实验在中国空气动力研究与发展中心(CARDC)Φ0.5 m高超声速风洞中进行,来流Mach数为5。实验模型为椭圆转圆形内转进气道,总收缩比为5.8,内部收缩比为1.7,喉部为直径50 mm的圆形截面。模型的肩部区域种植了长度与间隔可更换的丝线,为了改善进气道的起动性能,模型进气道的内压缩段开设了可以动态堵塞的泄流孔,在喉道下游设置了可动态节流的节流锥。实验获得了丝线长度、相邻丝线间隔的推荐值,同时表明,丝线流动显示技术能够快速、准确、直观、方便地判断进气道的起动状态,并能定量给出流动分离起始位置与分离结构,所采用的丝线流动显示技术丰富了高超声速风洞实验的流场可视化方法库。研究还表明,采用丝线流动显示技术,所研究的内转进气道在Ma=5时处于双解区,实验给出了进气道重起动及退出不起动的一种可行方案。

激波串与进气道肩部分离泡相互作用

针对高超声速进气道激波串与肩部分离泡相互作用时的流动振荡问题,在来流马赫数为6的激波风洞中,采用高速纹影拍摄结合壁面动态压力测量,研究了有/无抽吸情况下激波串与肩部分离泡的相互作用过程。结果表明:当激波串前移至肩部附近时,有抽吸进气道也会产生大尺度的分离泡,进而有/无抽吸进气道内的激波串均会与肩部分离泡形成耦合振荡,并造成严重的脉动压力。在激波串的推动下,分离泡能够自由地越过肩部凸拐角,使得其自身的低频振荡特性能够显现。激波串内的压力波动会显著改变分离泡的形态,而分离泡形态的变化又会影响激波串内的压力,两者相互耦合从而维持这种低频振荡。无抽吸进气道具有相同的低频耦合振荡特性;而抽吸缝阻碍了上下游的信息传递,使得有抽吸进气道的分离泡低频振荡显著,而激波串振荡具有一定的宽频特性。经分离激波振荡范围和进气道入口速度无量纲后,有/无抽吸进气道低频耦合振荡的St均处于0.011~0.021,与经典分离泡的低频振荡特性相当。

Design concept of three-dimensional section controllable internal waverider hypersonic inlet

A new hypersonic inlet named three-dimensional section controllable internal waverider inlet is presented in this paper to achieve the goal of section shape geometric transition and complete capture of the upstream mass. On the basis of the association between hypersonic waverider airframe and streamtraced hypersonic inlet, the waverider concept is extended to yield results for the internal flows, namely internal waverider concept. It is proven theoretically that not osculating cones but osculating axisymmetric theory is appropriate for the design of section controllable internal waverider inlet. And two design methods out of the internal waverider concept are proposed subsequently to construct two inlets with specific section shape request, triangle to ellipse and rectangle to ellipse ones. The calculation results show that the inlets are capable of keeping their shock structures and the main flow characteristics exactly as their derived flowfield. Further, the inlets successfully capture all the upstream mass despite their complicated cross-section transitions. It is believed that the concept proposed ex- plores a new way of designing three-dimensional hypersonic inlets with special demand of section shape transition. However, the detailed flow characteristic and the performance of the internal waverider inlets are still under investigation.