A model of a NUAA-PTRE pre-cooled air turbine engine was established. The design point parameters of the engine were optimized, including the pressure ratio, air flow rate of the compressor, efficiency, throat area, and efficiency of the turbine. The air flow rate at the engine operating point was 142.73 kg/s. High performance of the key components under a wide range of working conditions was realized after optimization. To achieve the indicators of the overall scheme, adaptability studies of key components were conducted. A three-stage variable geometry design was applied to the inlet. The pre-cooler was optimized with a power-to-weight ratio of over 100 kW/kg and a compactness of 278 m2/m3. The built-in rocket gas generator and dual-component injector were developed, and the combustion and heat transfer processes were simulated. The overall optimization design of the NUAA-PTRE and the adaptive design of the components were completed, and high performance of the engine in a wide range of flight conditions at Ma 0~5 and altitude 0~25 km was achieved.

Fighters have increasingly demanding performance on both aerodynamic and stealth. To adapt the design requirement of various irregular-shaped supersonic inlets integrated with flat fuselages, a design method for non-axisymmetric generalized internal conical (NAGIC) flowfield is discussed and then developed. By osculating axisymmetric flow concept and method of characteristics, the developed supersonic NAGIC flowfield can be inversely designed with arbitrary capture curves, controllable incident/terminal shock wave geometries, and post terminal shock wave pressure. Employing the streamline tracing technique, a set of external-compression inward-turning inlets with triangle-like flat capture shapes for supersonic stealth fighters are obtained to verify the design method. The study find that the internal contraction ratio and flow distortion of the inlet are reduced significantly by replacing the weak reflected shock wave in the original hypersonic NAGIC flowfield with a strong one. Thus, the compatibility between inlet and engine is improved in a relatively wide range. In addition, the irregular-shaped curved-compression inlet has high mass ratio and total pressure recovery ratio with negligible cross flow and uniform flow at throat. The performance of inlets with complex configurations can be further improved by matching the entrance geometries of basic flowfields and inlets.

以双通道涡轮/冲压组合发动机为例,在不影响涡轮通道工作状况条件下,研究了低速段冲压通道富余气流处理方式对发动机阻力特性影响.采用动量定理与壁面积分两种方法对比了冲压通道开启与关闭两种方案下发动机总体受力情况,提出了一种冲压通道冷态尾喷管节流减阻措施.研究表明:两种受力分析方法结果一致,均能准确评估发动机受力;相比于冲压通道关闭、气流高速绕过发动机的情况,冲压通道开启状态下,进气道溢流阻力小,即富余高速气流经冲压通道减速后,直接排出发动机所产生的阻力更低,且来流马赫数越低,两种方案的阻力差距越大;此外,合理控制冲压通道下游尾喷管喉道面积,降低冲压通道内的流速对发动机总体有明显减阻效果,Ma∞=0.8、1.0时减阻最为明显,最高减阻54.0％.

The inverse design of super/hypersonic flows is widely utilized in aerospace, especially in waveriders, inlets, and nozzles. However, most of the existing methods are intended for the two-dimensional (2D) problem. The inverse method for generalized three-dimensional (3D) supersonic flows is still immature and is the main purpose of the space-streamline-based method of characteristics (SMOC) presented in this paper. The key of SMOC is to integrate an additional Euler predictor–corrector algorithm for pressure gradients in the unit calculation process. In this way, the temporary orientation of the osculating plane (OP) of the space streamline is determined, and the conventional 2D axisymmetric method of characteristics can be adopted in the OP. Three common unit processes of SMOC and the posedness are introduced, and the astringency is demonstrated by corresponding algebraic calculations. With this method, inviscid super/hypersonic flows can be solved on the basis of specified flow features, such as a 3D shock surface or a 3D wall pressure distribution. The accuracy and efficiency of SMOC are verified by using an inverse design example, that is, the flow produced by an elliptic conical surface at a freestream Mach number of 6. The numerical simulation of the inverse design result indicates that the 3D shock wave geometry and the 3D wall pressure distribution match the targets completely. The relative root-mean-squared error of the surface geometry is 10−3 magnitude, and the computation time cost of the inverse design is less than that of the general direct Euler solver.

The inverse design of super/hypersonic flows is widely utilized in aerospace, especially in waveriders, inlets, and nozzles. However, most of the existing methods are intended for the two-dimensional (2D) problem. The inverse method for generalized three-dimensional (3D) supersonic flows is still immature and is the main purpose of the space-streamline-based method of characteristics (SMOC) presented in this paper. The key of SMOC is to integrate an additional Euler predictor-corrector algorithm for pressure gradients in the unit calculation process. In this way, the temporary orientation of the osculating plane (OP) of space streamline is determined, and the conventional 2D axisymmetric method of characteristics can be adopted in the OP. Three common unit processes of SMOC and the posedness are introduced, and the astringency is demonstrated by corresponding algebraic calculations. With this method, inviscid super/hypersonic flows can be solved on the basis of specified flow features, such as a 3D shock surface or a 3D wall pressure distribution. The accuracy and efficiency of SMOC are verified by using an inverse design example, that is, the flow produced by an elliptic conical surface at a freestream Mach number of 6. The numerical simulation of the inverse design result indicates that the 3D shock wave geometry and the 3D wall pressure distribution match the targets completely. The relative root mean squared error of the surface geometry is 10−3 magnitude, and the computation time cost of the inverse design is less than that of the general direct Euler solver.

Conventional compression profile of a chin inlet has an inevitable 3D effect, particularly under the condition of large angle of attack. This condition results in overflow and decrease of mass capture coefficient. In response to this problem, based on the variable geometry adjustment scheme of the original throat slider moving forward and backward, an integrated design of the chin inlet with waverider forebody and practical modification is conducted. Numerical studies show that the waverider forebody variable geometry chin inlet is designed with a working range of Ma 2.5 to Ma 4.0, the lift-drag ratio is 2.54 at the design condition, and the mass capture coefficient is 1.140, which is 3.6% higher than that of the original conventional inlet. Furthermore, the exit critical total pressure recovery coefficient and pressure ratio are 0.501 and 74.54, which are 9.6% and 11.8% higher than the original inlet, respectively. The practical modification can significantly reduce the size of the forebody. Although the performance of the inlet is slightly lower than before the modification, it is still higher than that of the original conventional inlet.

采用数值模拟的方法研究了航空发动机微引射式防冰腔结构参数对防冰热效率和发动机蒙皮温度分布的影响.通过改变射流孔直径、射流孔孔距、混合腔长径比和波纹板通道出口高度,建立了不同的微引射式防冰腔模型.在进口热气流量一致的情况下,对不同防冰腔模型进行了数值计算.结果表明:减小射流孔直径和波纹板通道出口截面高度能提高防冰热效率和蒙皮表面平均温度;射流孔孔距和混合腔长径比存在一个最优值使防冰腔换热性能最好.

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.

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.

传统的密切轴对称理论被广泛应用于均匀来流下的三维密切曲面激波反设计,为解决非均匀来流条件下的三维曲面激波反问题,提出了一种微元密切轴对称流场(MOA)求解方法.该方法沿激波面的周向和流向构建一系列微元密切面,在每个微元面内进行三维向二维流动的等效转换,从而突破了传统密切方法中不能有横向波后流动的限制.利用该方法编写设计程序,分别基于带攻角来流条件和外锥型流来流条件重构了标准内锥曲面激波,并与数值仿真结果进行了比较,结果表明,非均匀来流下激波曲面的三维形状均与预设形状完全一致,实现了非均匀来流下曲面激波形状可控.MOA方法在吸气式高超声速推进领域中前体∕进气道一体化设计方面有重要应用前景.

To provide ease of integration design of forebody/hypersonic inward turning inlet as well as to extend the domain of option for basic flowfield of the inlet, a design method for axisymmetric basic flowfield with controllable incident/reflected shock wave and certain post reflected shock parameter distribution is developed by using the method of characteristics and optimization algorithm. On this basis, a design method for non-axisymmetric generalized internal conical (NAGIC) basic flowfield is proposed by using the theory of osculating axisymmetric flows to incident and reflected shock wave in parallel. Numerical simulations are performed to verify the design methods. Detailed flow characteristics of NAGIC flowfield with an elliptical entrance that is assigned somewhat arbitrarily are presented. The incident and reflected shock wave geometries are controllable as given. In cross sections, the shock wave geometries are similar to the entrance shape of the inner cone. In osculating planes, the incident and reflected shock waves attach accurately to the lip and the shoulder, respectively. In addition, the crossflow in the 3D flowfield is negligible, and the flow parameters in the exit plane are almost uniform. Thus, the NAGIC flowfield is suitable for inward turning inlet design with complex configurations.

Traditional osculating axisymmetric flow (OA) theory is widely used in the inverse design of generalized shock waves. However, this theory is only applicable to uniform incoming flows. In this paper, a novel method called micro osculating axisymmetric flow (MOA) method is proposed to solve the inverse problem of the 3D generalized shock wave design under non-uniform incoming flows. This method constructs a series of micro osculating planes (MOPs) along the shock wave surface in spanwise and streamwise directions. Actual 3D flows are then approximated by 2D axisymmetric flows in each MOP. Thus, the new method breaks the restriction of no lateral velocities or pressure gradients in the traditional method. An internal conical shock wave at a 4° angle of attack and the other one in an external conical flow of a 10° cone half-angle are obtained by the novel method to validate its feasibility and applicability. Numerical simulation results of the two cases indicate that the 3D shock wave geometries completely match the target. A streamwise integration of the inward turning wavecatcher inlet and outward turning waverider forebody is also presented. This integration shows the remarkable application prospects of the MOA method in the field of air-breathing hypersonic propulsion.

针对马赫数0～6的预冷涡轮+冲压组合多热力循环发动机的宽范围工作要求,提出了一种在马赫数2～6范围内流量系数为1.0的宽范围轴对称进气道变几何调节方案,通过中心锥与分流板的协同平移运动,可在满足涡轮与冲压两通道流量分配要求的同时,实现两通道压缩量的匹配调节.对起始半锥角分别为20°和13°的两种变几何进气道方案开展了设计与对比研究.结果 表明:起始半锥角对最终方案设计影响最大,起始半锥角为13°的进气道方案较起始半锥角为20°的方案,冲压通道和涡轮通道在来流马赫数为6时临界总压恢复分别提高了16％和14％,最大迎风面积减小了12.4％,但中心锥和分流板平移调节距离分别增加84％和91％.

针对某固冲发动机用“X”型布局超声速混压式进气道展开进气道/弹体一体化数值仿真研究和吹风实验,研究结果表明:进气道的性能参数受到来流马赫数(Ma)和攻角(α)的影响,由此获得的拟合公式可预估不同来流条件的气动性能;“X”型布局的气动规律表现为性能参数随马赫数呈线性下降,随着攻角呈3次曲线下降趋势,实验结果随攻角下降的幅度更大;同时发现标准k-ε模型在大攻角(α=7°)下模拟精度不及其他湍流模型,而在更大攻角(α=10°)下RANS模型对流场分离的预测效果欠佳.

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%。

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。

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.

为研究内转式进气道进出口宽高比对流动特性的影响规律,通过设置宽高比不等的矩形进出口形状,对不同进出口宽高比组合方式下的进气道性能进行了数值模拟.在Ma6设计点对多个模型的计算结果表明:出口与进口宽高比的比值是影响进气道性能的关键参数,该比值在1～1.5内进气道总压恢复系数较高,最大增幅达25.3％;该比值为1时进气道出口气流畸变最大,通过设置合理的进出口比值可有效减小出口气流畸变,且最大降幅达40.5％;此外,进出口宽高比在0.8～2.4内,增大进出口宽高比可最多缩短25％的进气道长度.

To explore performance potential of hypersonic curved shock compression, high performance curved shock compression design results were obtained through multi-objective optimization and sampling com-putation, and were compared with traditional compression, under both inviscid and viscous condition. The re-sults show that,(1) under inviscid condition,in the Pareto solutions of Multi-objective optimization,there ex-ist curved shock compression cases,which have shorter length,higher compression efficiency and higher mass capture ratio under low flight Mach number, than traditional equal shock strength three ramps compression;(2) under viscous condition,when pressure ratio is 8.29 (1±0.5%),compared with isentropic compression, the length of curved compression is 35%shorter;compared with traditional equal shock strength three ramps com-pression,the length of curved compression is 8%shorter,the total pressure ratio is approximately equal,and the mass capture ratio under flight Mach number of 4 is 6% higher; compared with ramp-isentropic compres-sion, the total pressure ratio of curved compression is 1.1% lower, while the mass capture ratio under flight Mach number of 4 is 6% higher; outlet boundary layer thickness and shape factor of curved shock compression are lower than those of other compression methods.

In order to improve the seal and aerodynamic heating at the variable geometry throat section of RBCC inlet at high Mach number,an over/under type inlet was discussed by referring to the corresponding inlet that was previously applied to the TBCC.In the range of low Mach number,the inlet was adjusted to get a prop-er total contraction ratio and the same internal contraction ratio of each passage. Thus the two passages could gain similar performance parameters to reduce the mixing loss and resist the roughly same back pressure.In the range of high Mach number,the low-speed passage was closed and the high-speed passage was kept fixed.The numerical simulation results indicate when the free stream Mach number are 2.0, 2.5, 3.0, 3.5, the average Mach number of the two throat and outlet sections are similar, the deviations do not exceed 0.05 and the mass flow rate reaches about 0.7.In summary,this design enables the inlet to possess a high mass flow rate and good performance during the whole working range Ma=2~6,and the adjusting method is practical and feasible.