针对传统水平排气劈缝结构流动损失大的问题,提出了径向倾斜排气的设计思想,并围绕其进一步提出了直线式和曲线式两种倾斜劈缝结构.基于数值仿真研究了其内部流场,揭示了倾斜劈缝可减小冷气转折角,抑制旋涡产生,使流动更加平缓,从而减小流阻的机理.通过与水平劈缝的对比分析,初步验证了两种新型劈缝结构使总压损失分别降低了约10％～12％和13％～15％.进一步考虑了燃气外流与冷却气掺混过程对劈缝内流动的影响,仿真结果同样印证了倾斜劈缝对于内流减阻能力的提高.但同时也发现了此类倾斜射流导致掺混损失增大的现象,通过给出两类劈缝结构的涡轮叶栅整体流动损失变化规律为叶片综合性能优化提供了参考.

Compressed air energy storage systems must promptly adapt to power network demand fluctuations, necessitating a high surge margin in the compression system to ensure safety. It is challenging to completely eliminate blade geometric variations caused by limited machining precision, the important effects of which should be considered during aerodynamic shape design and production inspection. The present paper explores the uncertainty impact of geometric deviations on the stability margin of a multi-stage axial compressor at a low rotational speed. Initially, an adaptive polynomial chaos expansion-based universal Kriging model is introduced, and its superior response performance in addressing high-dimensional uncertainty quantification problems is validated through rigorous analytical and engineering tests. Then, this model is used to statistically evaluate the stability margin improvement (SMI) of the compressor due to the Gaussian and realistic geometric variabilities separately. The results show that the mean and standard deviation of SMI are −0.11% and 0.5% under the Gaussian geometric variability, while those are 0.33% and 0.39% under the realistic variability. For both the geometric variabilities, the stagger angle and maximum thickness deviations of the first-stage rotor are the most influential parameters controlling the uncertainty variations in the stability margin. Finally, the underlying impact mechanism of the influential geometric deviations is investigated. The variation in the stability margin caused by the geometric deviations primarily results from the alteration of inlet incidences, affecting the size of the tip leakage vortex blockage and boundary-layer separation regions near the blade tip of the first-stage rotor.

In this paper, an adaptive sparse arbitrary polynomial chaos expansion (PCE) is first proposed to quantify the performance impact of realistic multi-dimensional manufacturing uncertainties. The Stieltjes algorithm is employed to generate the PCE basis functions concerning geometric variations with arbitrary distributions. The basis-adaptive Bayesian compressive sensing algorithm is introduced to retain a small number of significant PCE basis functions, requiring fewer model training samples while preserving fitting accuracy. Second, several benchmark tests are used to verify the computational efficiency and accuracy of the proposed method. Eventually, the coexistence effects of six typical machining deviations on the aerodynamic performance and flow fields of a controlled diffusion compressor cascade are investigated. The probability distributions of the machining deviations are approximated by limited measurement data using kernel density estimation. By uncertainty quantification, it can be learned that the mean performance seriously deteriorates with increasing incidences, while the performance at negative incidences is more dispersed. By global sensitivity analysis, the leading-edge profile error should be given high priority when working at negative incidences, and the inlet metal angle error would be carefully inspected first when the cascade works at high positive incidences. Furthermore, controlling the manufacturing accuracy of the suction surface profile error can play a certain role in improving the robustness of aerodynamic performance in off-design conditions. Through flow field analysis, it further proves that actual leading-edge errors are the most important ones to aerodynamics and reveals how the effects of leading-edge errors propagate in the cascade passage, thus affecting the aerodynamic loss.

The impact of geometric deviation due to manufacturing on compressor performance is considerable in engineering practice. To investigate the impact of blade thickness deviation on compressor performance and flow loss at various rotational speeds, a three-dimensional steady numerical simulation on Rotor 37 was conducted. The quantification of uncertainty was accomplished using a non-intrusive polynomial chaos method. The viscous dissipation coefficient was introduced to analyze the uncertain influence of blade thickness deviation on flow loss. Based on the type of loss source, the flow field was divided into six regions, including the blade tip region, blade root region, leading edge region, trailing edge region, blade surface region, and mainstream region. The results indicate that the sensitivity of total pressure ratio to thickness deviation increases significantly with an increase in the rotational speed. Under peak efficiency conditions, the effect of blade thickness deviation on flow dissipation in leading edge region decreases initially and then increases with an increase in the rotational speed. Meanwhile, the impact on flow loss in other regions increases with the increase in the rotational speed. Under near stall conditions, the blade thickness deviation has a great impact on the flow losses in the blade tip region, leading edge region, and mainstream region at 60% design rotational speed. However, the blade tip region and trailing edge region are more noticeably affected at 100% design rotational speed. Furthermore, the quantification of standard deviation of flow losses in various regions under different rotational speeds and conditions reveals that the flow loss fluctuation in the leading edge region and mainstream region varies with changes in operating conditions and rotational speeds, but the fluctuation of flow loss in other regions is independent of the rotational speed.

基于双外涵变循环发动机压缩系统,分析了多连通气动布局变循环压缩系统的匹配工作机制.一体化全三维数值模拟表明:变循环压缩系统各压缩部件与涵道及其调节机构之间由于多连通特征相对于常规发动机压缩系统具有更强的耦合工作特点,高效的外涵道流动是发挥变循环发动机性能优势的关键.涵道几何的调节不仅会改变其自身流动状态,还伴随着压缩部件气动性能的偏移,模式转换过程必须符合各涵道及调节机构之间的气动协调匹配.提出了适用于多连通变循环压缩系统的一体化变维度分析方法,将部件通流程序与涵道零维程序相结合,实现了部件-涵道耦合匹配关系的快速分析.基于变维度分析方法给出了单外涵模式部件与涵道共同约束下的压缩系统综合匹配可行域,旨在为变循环发动机的匹配设计提供理论依据.

为了研究双外涵变循环压缩系统的涵道流动匹配规律,发展了涵道流动的计算模型,对压缩系统模式转换过程中外涵道倒流问题的发生机理和影响因素进行了深入分析,总结了压缩系统的倒流判断准则,并基于简化计算模型和全三维计算结果对倒流判据进行了验证.研究结果表明,对于一定的压缩系统匹配状态,存在一个决定涵道匹配状态的临界核心机驱动风扇级(Core Driven Fan Stage,CDFS)总压比,当CDFS的实际工作压比高于临界压比时,压缩系统第二外涵道将发生倒流,反之则系统不倒流.复杂涵道系统内的流动损失和堵塞等特性对临界CDFS总压比有显著影响,为了准确判断压缩系统的匹配状态,需要对其进行精确模化.基于简化模型得到的压缩系统倒流临界线可以推广至全三维状态,提出的倒流判断准则具有较高的可靠性和可行性.

随着飞行器对航空发动机的能力需求和发动机自身技术难度的提高,其研制的风险、周期、成本不断上升,对采办模式也提出了更高的要求.提出将发动机采办过程数字化的思路,从联合需求分析入手,运用多维缩放技术,以虚拟样机各维度的仿真模型为基础,开展一体化仿真,实现对技术方案的评估;以工作分解结构为纽带,建立技术方案、进度和费用等的映射关系,构建发动机虚拟采办的基本框架,实现物理域模型与业务域模型的有效集成,可快速预测进度和费用,提升采办评估的科学性.

Based on the commercial software Isight,a steady-state performance optimization method was proposed for a triple bypass variable cycle engine,and the optimization design platform was built.In this way,the function of the steady-state performance procedures was expanded,especially in the parameter study and optimization.Utilizing the platform,performance optimization of the typical engine working points was conducted,and the results were compared with the traditional method.With the help of the platform,the engine steady-state performance in different working conditions and various modes was analysed.The study results show that the Isight based optimization design platform could obviously improve the design efficiency and optimization precision,and fully exert performance advantage of triple bypass variable cycle engine.For getting the maximal thrust in any conditions,the best order of the mode selection is as follows as the inlet temperature raises:the first is single-bypass adding third-bypass mode,the second is single-bypass mode,the third is triple bypass mode and the last is double-bypass mode.

Aim to provide references for fighter engine technology investigation and product development, the development course and results of U.S. fighter engine were summarized based on the literatures. The development characteristics and tendency of U.S. fighter engine were summarized based on results of introducing fighter engine development. The research results show that technical development is solid foundation for product development, core engine and engine demonstrator are used for technical development and demonstration, and the adaptive cycle technology need to be demonstrated deeply. In terms of product development the engine model development pulled by national defense weapon demand was organically integrated by the various advanced technologies. The development tendency of the fighter engine is the engine performance improvement all the time, and the improvement of the engine reliability and durability is the inevitable requirement, and that would become the development tendency. The final goal of aero engine development is the economic affordability improvement.

分析了国内外飞/发性能一体化设计技术的发展,中国现处于飞机和发动机双方通过协调确定各自技术状态阶段,尚未开展系统一体化设计.详细阐述了前机身、进气道与发动机在流场、流量及隐身性能,发动机尾喷管与飞机后机身在安装性能、隐身性能,飞机功率提取、环控引气与飞/发性能、稳定性等一体化设计技术的研究内容.详细分析了在航空发动机研制中应用飞/发性能一体化技术的主要关注点,并指出了飞/发一体化设计技术推广应用的方向.

A new testing research on an existing engine provides an efficient way to explore critical engine technology.A new method is proposed to develop VCE(variable cycle engine) from a conventional gas turbine(turbofan) engine.This paper also presents a configuration promotion project with introduction of selector valve,and puts forward VABI(variable area bypass injector) and rear VABI.The first stage of high pressure compressor is changed to a CDFS(core-driven fan stage).The working principle of the VCE is stated.Based on the components′ performance level of conventional turbofan engine,a primary overall performance project is carried out and a numerical simulation on VCE is conducted.The calculation results show that the thrust can be increased by 5% to 8% in supersonic cruise using single bypass mode.Although specific fuel consumption(SFC) cannot be decreased obviously in subsonic cruise using double bypass mode,the engine inlet air flux can be increased by 5%,so the spillover is reduced and installed performance of the engine can be improved.

During the starting process of turbofan engine,the optimization of fuel supply in main combustor is an important key technology of successful start.According to the components characteristics of main combustor,and considering the accuracy of the engine control system,this paper puts forward a design method of asymptotic ignition for fuel supply control law.The method mainly includes the basic fuel quantity,the rate of rise and maximum of ignition fuel supply.Based on this,the judgment conditions of successful ignition are presented.It is applied to a certain type of turbofan engine.With the analysis of ground test,simulated altitude test and test flight,the method is proved to be more practical and effective.

It is an efficient way to study a new technical rearch on an exist engine. Based on a general gas turbine (turbofan) engine, a new method of verification is proposed for promoting variable cycle engine (VCE). This paper also presents a configuration promotion project with introduction of selector valve, forward variable area bypass injector (VABI) and rear VABI. The first stage of high pressure compressor is change to a core-driven fan stage (CDFS). The working principle of the VCE is stated. Based on the components performance level of general turbofan engine, a primary performance project is made and a numerical simulation on VCE is conducted. The calculation results show that the thrust can be increased by 5% to 8% in supersonic cruise using single bypass mode. Although specific fuel consumption cannot be decreased obviously in subsonic cruise using double bypass mode, the engine inlet air flux can be increased by 5%. And so the spillover is reduced and the intalled engine performance could be improved.

According to the basic theory of aerodynamic stability and engineering experience,an aero-engine,for example,aerodynamic stability design method in the whole flight envelope was presented on the basis of GJB/Z 224-2005,and some experimental verification were performed.The results show that stability margin required of fan and compressor is larger at high-altitude and low-speed state,stability margin required of fan is larger at right boundary of the engine envelope and basically remains the same,affected by the nozzle subcritical,stability margin required of fan is larger during landing.Therefore,these conditions should be focused in the engine design and use.Engine should be designed to ensure that the stability margin available of fan and compressor is greater than the stability margin required.The design method can effectively meet the demand for complex engineering development.