【作者】 李宇飞；

【导师】 何国强；

【作者基本信息】 西北工业大学 ， 航空宇航推进理论与工程， 2004， 硕士

【摘要】 火箭基组合循环发动机(RBCC：Rocket Based Combined Cycle)是单级入轨最有潜力的动力方案之一，具有广泛的军事和民用航天应用前景，受到世界各国航天界的瞩目。引射火箭模态用于RBCC发动机起飞低速阶段，对于实现RBCC结构一体化、减轻结构重量和系统复杂度具有重要的意义。本文针对RBCC引射火箭的结构一体化和性能增强问题，对其进行了理论、数值计算以及实验研究。 在理论研究方面，本文对引射火箭二次加热及热力喉道进行了研究，推导并分析了一般的管流理论，阐述了加热管流与变截面等熵管流之间的异同。结果认为，对于后体为纯扩张构型的引射火箭，不论发动机通道内的马赫数是多少，都可以通过加热(改变加热位置)代替常规几何构型所起的作用。本文还从发动机受力角度以及热力循环角度，阐述了发动机各个部件对于发动机性能的影响，说明热力喉道可以取代几何喉道，以便使引射火箭构型适应超燃冲压发动机的构型。另外还得到热力喉道取代几何喉道后对发动机性能的影响，它可以提高发动机的推力，却要降低比冲。 在数值计算方面，本文采用了新的计算域，减小了由于边界条件造成的计算误差。通过数值计算研究了引射火箭堵塞比、来流马赫数、一次火箭燃烧室总压、一次火箭构型对引射火箭性能的影响。由数值计算获得的结论为：引射火箭内的一般流态为超音速区与亚音速区共存，亚音速区不具备做功能力；飞行马赫数增大，引射火箭内更易于产生壅塞；为了获得最佳的性能，随着飞行马赫数的提高，应当降低一次火箭的工作压强。 在实验研究方面，本文改进并利用现有的实验设施，进行了以酒精作为二次燃料的引射火箭二次燃烧实验。实验研究了二次燃料喷注位置、二次燃料喷注量、后体出口面积对引射火箭性能的影响，得到的结论为：通过燃烧组织，引射火箭模态可以和其他工作模态(亚燃／超燃)共用燃烧室而不需做大的几何变化；二次燃料在引射火箭通道截面上的分布对于引射火箭的稳定工作有重要意义。分析认为，将喷注孔开在顶板的中央或采用SPI燃烧组织方式，使燃料从主流散播出去可以使引射火箭工作更加平稳。 本文认为，引射火箭推力增强可以通过对超音速的主流进行加热来实现，引射火箭结构设计以及参数匹配可以根据这一条来进行；引射火箭二次加热应遵循先加热超音速燃气流，后加热亚音速二次流的原则。采用SPI燃烧组织方式，可以实现上述原则，改善推力特性。更多还原

【Abstract】 Rocket Based Combined Cycle (RBCC) will be one of the promising propulsion systems for single stage to orbit vehicle, and has extensively application foreground in both military and civil usage. Ejectorjet is the alternative power of RBCC during low-speed operation. The utilization of ejectorjet will be important for structural integration, and will decrease weight and complexity of the engine. Theoretical, numerical computation and experimental approaches are used in this paper to investigate the problem of structure integration and performance strengthen on ejectorjet.In the theoretical aspect, secondary combustion in ejectorjet and thermal throat was researched. Theory of duct flow was conducted, the difference and homology between heat added flow and entropy fixed flow was described. In divergent duct, adding heat can substitute the usage of geometry structure whatever mach number is. From force and thermodynamic cycle aspects, the effect of every components on engine performance was discussed. It indicates that thermal throat can substitute geometry throat, so structure integration can be achieved in ejectrojet mode. In addition, it showed that thermal throat will increase engine thrust and decrease specific impulse.In numerical computation aspect, by implying new computational geometric field, computation error caused by boundary condition was decreased. The effect of area ratio, Mach number, total pressure of primary rocket and geometry structure of primary rocket on ejectorjet performance was studied. The main results are: supersonic region and subsonic region can co-exist in flow path simultaneously, subsonic flow is insignificance for engine performance; total pressure of primary rocket should be decreased when Mach number is increased.In the experimental aspect, some components of ejectorjet experimental system were improved, ejectorjet experiments using alcohol as secondary fuel were conducted. The effect of secondary fuel nozzle location, the mass flow rate of secondary fuel and outlet area of ejectorjet were studied. The main results are: the geometry structure of ejectorjet can be compatible with other RBCC operation modes by secondary combustion organization; the distribution of secondary fuel on ejectorjet cross section is significant forsecondary combustion stability. SPI combustion organization method is promising.Thrust augmentation of ejectorjet can be obtained by adding heat in supersonic main flow. Structure design and operation parameter adjust should follow this rule. Secondary combustion should mainly heat the supersonic flow. SPI combustion organization has the latent to improve ejectorjet performance.更多还原