The flow compression process design is critical for enhancing the efficiency and overall performance of hypersonic airbreathing propulsion. Many studies have sought to solve the challenging problems that inlet design presents, such as enhancing performance at off-design points, minimizing total pressure loss, and shortening the length. In this case, an inverse design approach for curved surfaces is proposed, which generates the entire compression flow field with a certain wall Mach number distribution as well as the molded line of the compression surface. In contrast to traditional flow compression modes, the proposed technique significantly increases design space and flexibility. Curved surfaces with a specific wall Mach number linear distribution along the flow direction are examined in this work. The flow compression process is made more efficient by using the inverse design approach. The curved surface compresses the flow primarily by isentropic compression and supplementary shock compression, and the entire wall surfaces participate in this process. In particular, the ratios of shock compression and isentropic compression can be changed as necessary. The result is a compromise between classical isentropic compression and multiramp compression. To illustrate the benefit, the incorporating of curved surface with sidewall compression inlet was evaluated and verified in wind tunnel tests. Mach number and pitot pressure profiles in the isolator's exit cross section and wall pressure distributions were utilized to study the flow field. The application provides a more flexible method and achieves the objective of functioning in a broad range of Mach numbers, according to the results. The experimental findings were found to be in accordance with the projected inlet performance.

Conventional forward bleed slots reduce the hypersonic inlet starting Mach number but suffer from excessive flow leakage after restart. This paper proposes a novel reverse bleed slot design method for curved axisymmetric inlets of a solid-fuel scramjet. Leveraging the Kantrowitz criterion and detailed flow analysis, the method optimizes bleed slot placement, number, area, and angle. Results show superior aerodynamic performance by placing slots in the non-starting region of the internal compression section, considering both unstarted flow and separation bubble dynamics during restart. Each bleed slot area is calculated successively down-stream based on the Kantrowitz criterion. Finally, the effects of bleed slot angle have been extensively studied. The key inlet performance reaches its optimum at a slot angle of approximately 130°, achieving a significant reduction in the starting Mach number (from 4.80 to 3.65) and a 50% decrease in bleed flow rate compared to the forward slot design. This method demonstrates its feasibility and effectiveness, enabling substantial improvement in inlet starting performance with minimal flow loss.

A new idea of nosebleed air jets with strong coupled internal and external flow is put forward using the lateral jet control principle to improve the maneuverability and fast reaction capabilities of hypersonic vehicles. The hypersonic vehicle’s nose stagnant high-pressure and high-temperature gas is utilized as the drive source for long-term jet control. The significant coupled jet interaction of the internal and external flow changes the aerodynamic characteristics. As a result, the structure is basic and does not rely on any external source to achieve flight attitude control. The complicated flow characteristics of the nosebleed jet in supersonic crossflow surrounding the vehicle were numerically and experimentally investigated. The jet interaction characteristics and the aerodynamic characteristic changes generated by the nosebleed air jet are verified by comparing the flow field with and without the jet. Results indicate that the nosebleed air jet alters the center-of-pressure coefficient, which is subsequently coupled with the interference aerodynamic force. This results in a variation in pitch moment. The jet decreases the pitching moment coefficient when compared with the case without a jet. It is probable that combining nosebleed air jets with model centroid adjustment yields an optimal trim angle of attack.

为了满足腹部进气布局高超声速飞行器乘波前体与进气道一体化设计要求,发展了一种来流非均匀的马赫数分布可控内收缩进气道设计方法.在来流马赫数和壁面马赫数分布规律同时给定的前提下,通过有旋特征线法反设计轴对称基准流场,然后结合流线追踪技术生成圆形进口内收缩进气道,同时与传统基于均匀来流设计的内收缩进气道进行对比.数值仿真结果表明:非均匀来流的基准流场结构与设计预期一致,可以实现对整个流场的马赫数分布控制,且其压缩效率高于传统均匀来流设计的基准流场.设计点时非均匀来流设计的进气道保持了基准流场的波系结构并实现了全流量捕获.有黏时非均匀来流设计的进气道总体性能较高且高于同样来流条件下均匀来流设计的进气道.该设计方法可行,为高超声速乘波前体与进气道一体化设计提供了一种新途径.

With the development of combined cycle engines, it is urgent to estimate more quickly and accurately the flow capture capacity and starting performance of variable geometry inlets over a wide Mach number range. Based on the flow field and parameter fitting, two prediction methods for the curved axisymmetric inlet with lip translation scheme have been proposed. The method based on the flow field of the reference inlet is more efficient than the parameters-based prediction method, as it can accurately predict the lip translation distance and the corresponding flow coefficient over the entire working range of the inlet without additional numerical calculations. Moreover, the starting Mach number is accurately predicted by the fitting method based on the throat Mach number of the reference inlet, with a relative error of only 0.95% compared to the numerical simulation. The flow coefficient-based method is simple and accurate for predicting lip translation distances with a known starting Mach number, with a relative error of only 1.65% compared to numerical simulations. The prediction approaches can overcome the drawbacks of the standard iterative algorithms and significantly enhance computational accuracy and efficiency.

采用内外压缩型面可控的弯曲压缩进气道反设计方法,设计了一种新型混合并联式二元弯曲压缩进气道,重点针对进气道模态转换过程及冲压单独工作条件下进行了数值仿真及试验研究,获得了进气道宽速域性能,结果也表明新型混合并联式弯曲压缩进气道具有较高的综合气动特性,冲压工作状态马赫数为4、攻角为3.时总压恢复0.5以上,马赫数为3、攻角为3.时总压恢复0.75以上,涡轮工作状态马赫数为2、攻角为3.时总压恢复0.88以上,综合畸变指数小于5％,满足宽速域进发匹配需求.

For scramjet, the inlet plays the role of capturing airflow, pressurizing and decelerating the airflow, and providing supersonic airflow to meet the combustion requirements for the downstream combustion chamber. In this paper, a prototype inlet is designed by using the method of spiral characteristic lines, and then the influence of moving center cone on the inlet performance is studied by numerical simulation. It can be found that the method of moving the center cone can obviously improve the performance of non-design point and self-starting performance of the inlet.

附面层边缘通常取在速度达到主流速度0.99倍的位置,而复杂流场中主流流动往往并不均匀,给附面层边缘的准确识别造成了困难.为解决此问题,提出了用“参考主流”代替实际主流识别附面层边缘的方法:通过零剪切力滑移壁面边界条件下数值模拟得到不受附面层干扰的参考主流,在根据附面层定义确定附面层边缘时以该参考主流中的速度代替实际的主流速度.通过斜楔压缩和弯曲压缩两个超声速压缩流场对该识别方法进行了验证,所得到的斜楔压缩出口截面上附面层厚度与采用实际主流速度判断得到的厚度相对误差仅4.1％.根据该方法的识别结果对弯曲压缩型面设计进行附面层修正后,弯曲激波高度与无黏设计值之间的误差从修正前的2.0％降低至0.3％,压缩面末端压力的相对误差从修正前的6.6％降低至2.3％.该方法避免了指定主流速度的主观性,识别结果较为准确.

针对宽范围定几何颌下进气道高马赫数下的压缩量不足问题,提出了一种喉部滑块前后移动的变几何调节方案,该方案通过滑块前后移动改变高低马赫数下的喉道尺寸,使进气道能够满足高低马赫数下的压缩量要求.提出了两种滑块布局方式,针对内锥侧滑块布局方式,按调节原理进行了滑块型面与进气道内流道型面的匹配设计,并将变几何颌下进气道与定几何方案进行了性能比较.数值研究表明:按Ma=2.5～4.0设计的变几何颌下进气道,在设计点,临界状态出口总压恢复系数为0.51,较公开文献中定几何方案提高8.5％;在Ma4.0,0°攻角工况下,临界状态出口总压恢复系数为0.46,提高12.2％;在Ma2.7,1°攻角工况下流量系数为0.69,临界状态出口总压恢复系数为0.78.气动性能表明,该颌下进气道性能优越,调节方案简单可行.

为提高五孔探针的标定效率,研究了一种采用拉丁超立方抽样技术开展变马赫数五孔探针标定的试验方法.马赫数在0.2～0.5范围内的对比试验结果表明:采用改进的拉丁超立方抽样标定技术可使攻角与侧滑角的拟合误差较一般拉丁超立方抽样方法下降14.3％～28.6％,且攻角与侧滑角的拟合方均根误差均在1°以内,马赫数的拟合方均根误差不高于0.005,标定精度能够满足工程使用要求.与常规完整数据标定方法相比,改进的拉丁超立方抽样标定技术虽拟合误差升高8.8％～17.1％,但标定工作量仅为常规完整标定方法的20％,标定成本下降80％,马赫数范围越宽,成本的降低越明显.

开展了内外压缩型面可控的弯曲压缩进气道反设计方法研究,并针对设计的二元弯曲压缩进气道进行了不同马赫数下变攻角起动数值仿真及试验研究,获得了进气道起动攻角迟滞环,仿真与试验得到的不起动/自起动攻角、出口流场以及压力分布吻合良好,同时试验结果也表明设计的弯曲压缩进气道具有较高的综合气动特性,设计点Ma=4、α=3°抗反压能力70倍来流压力以上,总压恢复系数在0.5以上.

为初步研究高马赫数内转进气道在真实气体效应下的工作特性,首先设计额定工作状态Ma=12的高超声速内转进气道,再结合不同气体模型对其进行数值模拟.研究结果表明:化学非平衡气体在流场结构、工作性能和气动加热方面与热完全气体较为相近,与热化学非平衡气体存在一定差别.离解反应发生在边界层内和低速涡流区内,热化学非平衡气体的离解反应程度比化学非平衡气体大.在隔离段内激波反射处,相比完全气体,化学反应气体的静温降低了2000～2500K.高热流区在上壁面喉道位置与下壁面激波反射点位置附近,温度较高的等温壁面、热化学非平衡气体均可降低壁面热流密度,不同壁面条件对隔离段出口性能参数影响较为明显.真实气体效应、壁面温度对隔离段涡流区的影响较为复杂,有待进一步研究.

基于对轴对称基准流场参数化研究选取半径适当小的可变中心体,再对其他设计参数进行灵敏度分析,得到设计参数对基准流场整体性能的影响规律,系数c的影响最为明显,同时各个设计参数之间耦合效应影响也很大.运用样本数据库,构建相应的神经网络近似模型并结合邻域培植多目标遗传算法对轴对称基准流场在马赫数为6时进行三目标优化,优化后的基准流场内收缩比降低了17.7％,总压恢复系数提高了2.3％,并且静压比提高了7.1％.基于此优化结果,进行内转进气道型面设计并对其在马赫数为3～6条件下黏性数值模拟,结果表明:优化后的内转进气道在马赫数为3工作时能够正常起动,在马赫数为4～6工作时,进气道有较高的压缩量,较好的流量捕获能力和总压恢复性能.

为了满足二元曲面可调进气道模态转换马赫数范围(来流马赫数为2.2～3.2)的流量要求,针对唇口平移、转动和转动+平移三种调节方案,基于理论分析和基准进气道的流场,提出了一种流量系数精确预测方法,并通过数值计算进行验证且获得了进气道的总体性能.结果 表明:调节后的进气道流量系数与预测值完全相等,而且无需多次试算,符合设计预期,可拓展应用于轴对称进气道.相对基准进气道,唇口前移时流量系数和压缩效率同时增加,来流马赫数为2.5时出口总压恢复系数相等而增压比增加了14.6％;在降低相同流量系数条件下,后移唇口使得增压比和压缩效率均降低,来流马赫数为2.5时出口总压恢复系数基本相等而增压比减小了12.9％,转动唇口使增压比进一步减小了9.1％,唇口后移方案性能更优.

为提高来流马赫数范围为2～4的"X"型进气系统大攻角下的稳定裕度,设计并研究了一种倒置二元进气道设计方案,并将其与正置方案进行了比较。结果表明:来流马赫数为2.3～3.5,攻角范围为0°～6°时,倒置布局设计方案总体性能较优,未出现明显激波/附面层干扰问题,能够满足设计要求。在采用相同的进气道设计方案时,倒置布局其迎风与背风进气道结尾激波位置及总体性能参数差异更小;0°攻角时倒置布局临界总压恢复系数与正置布局相当,4°攻角时倒置布局比正置布局高2%～3%,8°攻角时普遍高19%以上,且来流马赫数越高提升幅度越明显,8°攻角下倒置布局总流量系数较正置布局高6%左右。研究还发现,当来流马赫数较低时倒置布局总阻力低于正置布局,4°攻角时低1.7%;而来流马赫数较高时倒置布局总阻力高于正置布局,4°攻角时高2.0%。

侧偏捕获是高超声速内转进气道的常见捕获形式之一,但侧偏捕获对进气道起动的影响并不清楚.为探索侧偏捕获对高超声速内收缩进气道起动的影响,采用数值模拟方法研究了侧偏捕获进气道的起动机理.结果 表明,无量纲侧偏距离为0.04时,进气道再起动马赫数相对于对称进气道降低3.6％;无量纲侧偏距离为0.06时,进气道再起动马赫数降低7.6％.起动过程的边界层分离表明,内收缩进气道起动过程受分离区的影响很大.侧偏捕获进气道缩小了分离区,有利于再起动.相对于对称捕获进气道,侧偏捕获进气道会形成两个非对称、强度不等的流向涡.侧偏捕获减弱了流向涡对隔离段空间的占据作用,改善了内收缩进气道的性能.

针对超燃冲压发动机隔离段实际工作中,前方进气道唇口激波及肩部膨胀波对其性能的影响,采用真实入口条件隔离段简化模型,研究隔离段长度对隔离段性能的影响规律。结果表明:同一隔离段长度下,Ma4工况时真实入口条件隔离段模型的最大抗反压比均匀入口条件隔离段模型降低5.7%～8.5%,Ma6工况时降低5.9%～7%;获得相同最大压升下,Ma4、Ma6工况时真实入口条件隔离段长度分别比均匀入口条件隔离段长2H、H;从隔离段的抗反压能力角度考虑,真实入口条件下Ma4、Ma6工况的最佳隔离段长度分别为11H、13H,此时最大压升均可达到极限反压的96%以上。

Based on the rotational method of characteristics and streamline tracing tech-nique, the design method was developed for inward turning inlet with control of the axial and horizontal projections of intake, so as to meet the integrated design requirements of hyper-sonic vehicle waverider forebody and inlet with both sides intake layout.The inward turning inlet was designed utilizing this design method for the axial and horizontal projections of in-take of super ellipse.Numerical simulation was conducted at design point Mach number of 6.0.The results indicate that the axial and horizontal projections of intake conform to the anticipated design.The inlet can retain the wave structure and pressure distribution of basic flowfield, and capture all of free incoming flow under the inviscid condition.Its performance of throat plane is almost equal to basic flowfield.Under the viscous condition, the inlet has high compression efficiency;the flow coefficient and total pressure recovery coefficient of ex-it plane are 0.96 and 0.56, respectively.This design method is feasible and effective.

针对超燃冲压发动机隔离段实际工作中,前方进气道唇口激波及肩部膨胀波对其性能的影响,采用真实入口条件隔离段简化模型,研究隔离段内激波串迟滞回路现象.结果表明:均匀入口条件下升压与降压过程对隔离段内的流动特性基本不产生影响,而真实入口条件下升压和降压过程中隔离段内则出现了激波串迟滞回路现象;出现激波串迟滞回路现象时,同一反压下隔离段内激波串起始位置相差约为1.3倍隔离段高度,且激波串形态、壁面压升规律、总压恢复系数均差异明显,并得出附面层分离是导致出现激波串迟滞回路现象的主要原因.