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.

Numerical Investigation of an Active Jet Control Method for Hypersonic Inlet Restart

A flow control method based on an active jet is developed to restart hypersonic inlets. The dynamic restarting process is numerically reproduced by unsteady Reynolds averaged Navier-Stokes(RANS) modeling to verify the effectiveness and reveal the influence of jet conditions. The active jet improves the inlet unstart status by drawing the high-pressure separation bubble from the internal compression duct and performing a full expansion to alleviate the adverse pressure gradient. Moreover, the favorable pressure gradient in the inlet caused by jet expansion allows for a successful restart after turning off the jet. The influence of the jet momentum ratio is then analyzed to guide the design of the active jet control method and choose the proper momentum ratios. A low jet momentum does not eliminate the high-pressure separation bubble, whereas an excessive jet momentum causes severe momentum loss due to the induced shock. The general rule in restarting the inlet using an active jet is to allow a full jet expansion downstream of the jet slot while avoiding excessive momentum loss upstream and preventing the thick low-speed layer.

The Full Flowpath Analysis of a Hypersonic Vehicle

A numerical study has been carried out to investigate the full flow path and aerodynamic characteristics of a hypersonic vehicle at a 7.0 free stream Mach number. Results indicate that the inlet started and unstarted operations have remarkable effects on the flow pattern of the full flow path. When the inlet operates in a started mode, the transverse pressure gradient generated by the forebody alters the air captured characteristics and the entering flow quality of the inlet. Furthermore, the expansion process of the nozzle jet flow is obviously affected by the external flow field around the afterbody with the cross section shape transiting from a near rectangle at the exit of the nozzle to a near triangle at the tail of the vehicle. When the inlet operates in an unstarted mode, the aerodynamic instability can be observed in the full flow path of the vehicle. Due to the oscillation of the external compressed shock wave and nozzle jet flow, the aerodynamic characteristics of the vehicle vary periodically with the lift-drag ratio changing from 0.25 to 2.09. Finally, by comparing to the experimental data, the reliability of the CFD is verified.

Effect of expansion waves on cowl shock wave and boundary layer interaction in hypersonic inlet

The interaction of cowl shock wave and boundary layer has a crucial effect on the stability,operability and performance of hypersonic inlets.Many studies on inhibiting the sep-aration and managing the strength of the interaction of the shock wave and boundary layer with expansion corner have been conducted.However,the expansion waves near the circular arc shoulder to effectively control the interaction and cowl shock arrangement is little investigated.Therefore,the interaction of the cowl shock wave and boundary layer under thefluence of the expansion waves is studied by inviscid and viscous numerical simulations.The results reveal that the expansion waves have an important impact on the interaction between the cowl shock wave and boundary layer and the strength of shock wave,and that there are four types of inter-action processes with the change of the relative impingement positions of cowl shock wave.The expansion waves have a different influence on the shock wave and boundary layer inter-action at different incident points.When the incident point of the cowl shock wave goes far downstream from the end of the circular arc shoulder,the influence of expansion waves is weakened,and the magnitude of separation zone increases.However,when the expansion waves are applied to the interaction of the cowl shock wave and boundary layer on the circular arc shoulder,the separation can be effectively controlled.In particular,while the expansion waves interact with the shock wave and boundary layer in the back half of the circular arc shoulder,the separation is best inhibited.Compared with the upstream and downstream inci-dent points,the scale of separation area in the optimal control region is reduced by 65.3%at most.Furthermore,the total pressure recovery coefficientfirst increases and then decreases when the cowl moves from upstream to downstream,and the total pressure recovery coefficient reaches the maximum value of 68.36%at the incident position of cowl shock wave d Z 8.09d0.

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.

宽速域可调进气道密封腔流动特性研究

为了满足高超声速飞行器的宽速域进气需求,一般考虑使用可调进气道,此时进气道内部需要安装动密封结构。陶瓷栅片密封凭借其优异的耐磨损性和抗高温氧化能力是最理想的选择。使用陶瓷栅片作为密封件后,密封室内会形成受侧板边界层影响的三维空腔流动。以密封腔模型为基础,提出对密封腔流动和泄漏通道流动解耦化处理,并分别对进气道密封模型和泄漏模型进行了建模。利用数值模拟方法,改变来流边界层厚度以及来流速度研究不同来流条件对长密封腔结构类型流动特性及泄漏规律的影响。分析发现,长密封腔在Ma4.5高超声速来流下表现为开式空腔流动,对边界层厚度以及来流马赫数的变化相较于短密封腔更为敏感;来流马赫数的改变使得密封腔内流动结构变化较大,相应泄漏通道的入口压力分布也存在很大区别。

ANALYSIS OF INTERNAL WAVERIDER INLET AND TYPICAL SIDEWALL COMPRESSION INLET PERFORMANCE

A new internal waverider inlet with a rectangular shape of entrance and exit in front view is designed at Ma=6.0.The design is based on a better basic flowfield ICFC than traditional one and derived with the technology of stream tracing and shock cutting.Comparison between the newly designed inlet and a typical sidewall compression inlet is given.The design Mach number and entrance shape of this new inlet are chosen according to the sidewall compression inlet.Numerical results show that most of the performance parameters of the internal waverider inlet are a bit higher than the sidewall inlet,such as the flow capture coefficient,total pressure recovery and the kinetic efficiency.The performances of these two inlets at off-design points are compared.The internal waverider inlet can capture more than 91% of incoming flow under all simulated conditions.Results show that internal waverider inlet using 3-D compression and high flow capture coefficient is a kind of fixed-geometry inlet with better performance.

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。

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。

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。

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。

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.

飞行器前体-进气道热变形效应对进气性能影响的流/热/固耦合数值模拟

进气性能是吸气式高超声速飞行器进气道设计的重要指标。在长时间气动热载荷作用下,飞行器前体和进气道均会产生不同程度的热变形现象,改变进气道内部气流组织和流场结构,影响进气性能甚至危及飞行安全。基于自主研发的热环境/热响应耦合计算分析平台,开展了飞行器典型前体-进气道结构的流/热/固耦合数值模拟,分析了前体压缩面和唇口构型的热变形效应对进气道波系结构和进气性能的影响规律。分析表明:长时间巡航状态下,考虑热变形影响时,进气道唇口会偏离设计状态,波系随局部变形而发生位移和振荡,导致进气道入口流量系数上升,总压恢复系数下降,升压比上升。热变形导致进气性能相关影响需在吸气式飞行器设计中予以重点关注。

Design and analysis of a hypersonic inlet with an integrated bump/forebody

Numerical simulations and experiments showed that bump inlet had a remarkable effect on boundary layer diversion of supersonic flow.However,the design and analysis of bump in hypersonic flow was still few.In this paper,the mechanism of a supersonic bump inlet is introduced to the design of hypersonic forebody.A hypersonic inlet with an integrated bump/forebody is obtained by the Method Of Characteristics(MOC)based on a chin inlet.Numerical simulations show that the modified inlet achieves diversion of low-speed flow.Besides,the integrated bump/-forebody is also beneficial to inlet start.During the starting process,the shape of the separation zone is rebuilt by the modified forebody surface which makes spillage much easier.This new design leads to a reduction of the self-start Mach number by 0.95.

Numerical study of unsteady starting characteristics of a hypersonic inlet

The impulse and self starting characteristics of a mixed-compression hypersonic inlet designed at Mach number of 6.5 are studied by applying the unsteady computational fluid dynamics （CFD） method. The full Navier–Stokes equations are solved with the assumption of viscous perfect gas model, and the shear-stress transport （SST） k–x two-equation Reynolds averaged Navier– Stokes （RANS） model is used for turbulence modeling. Results indicate that during impulse starting, the flow field is divided into three zones with different aerodynamic parameters by primary shock and upstream-facing shock. The separation bubble on the shoulder of ramp undergoes a generating, growing, swallowing and disappearing process in sequence. But a separation bubble at the entrance of inlet exists until the freestream velocity is accelerated to the starting Mach number during self starting. The mass flux distribution of flow field is non-uniform because of the interaction between shock and boundary layer, so that the mass flow rate at throat is unsteady during impulse starting. The duration of impulse starting process increases almost linearly with the decrease of freestream Mach number but rises abruptly when the freestream Mach number approaches the starting Mach number. The accelerating performance of booster almost has no influence on the self starting ability of hypersonic inlet.