共查询到19条相似文献,搜索用时 125 毫秒
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为探究可重复使用火箭发动机设计参数对推力室身部在工作过程中热棘轮现象的影响,采用经典的液体火箭发动机设计方法设计了不同室压、推力及混合比的推力室,通过准二维传热计算方法、非线性有限元热-结构耦合分析方法和局部应变法对比了不同设计参数的推力室在工作过程中的棘轮应变及其发展情况。计算结果表明,相同的热试时间,循环工作的发动机推力室比单次工作的发动机推力室产生的应变更大;设计参数对棘轮应变的影响是通过改变推力室热环境来实现的;其他设计参数不变,室压更高、推力更小或混合比更高的推力室的棘轮应变更大;高室压、大推力或高混合比的推力室棘轮应变随循环次数的增加而减小。 相似文献
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一、前言目前,难熔金属在航宇工程上获得日益广泛的应用,它们在制造液体和固体火箭发动机推力室及喷管延伸段上显然占有重要地位,而为武器生产和航天事业所确认。宇航应用对制造火箭喷管的材料提出了特殊的要求。火箭和推力室喷管工作温度高, 相似文献
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为了在设计阶段提升带直二元喷管涡扇发动机的总体性能,本文开展了基于直二元喷管形面优化的涡扇发动机参数优化的研究。首先,建立了带直二元喷管的涡扇发动机模型,提出了发动机正后向排气系统红外辐射特征的计算方法,分析了直二元喷管尺寸对发动机性能参数的影响;其次,提出了基于序列二次规划算法的设计参数多目标优化方法,优化的目标包括高单位推力、低油耗和低红外辐射强度;最后,基于以上模型,利用序列二次规划算法对在设计点非加力情况下的涡扇发动机设计参数进行多目标优化。仿真结果表明:在设计点上,相较于不带直二元喷管的涡扇发动机,带直二元喷管的涡扇发动机具有更好的红外性能,并且通过算法优化后,带直二元喷管的涡扇发动机具有更好的性能参数。 相似文献
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为了在设计阶段提升带直二元喷管涡扇发动机的总体性能,开展了基于直二元喷管形面尺寸的涡扇发动机参数优化的研究。首先,建立了带直二元喷管的涡扇发动机模型,提出了发动机正后向排气系统红外辐射特征的计算方法,分析了直二元喷管尺寸对发动机性能参数的影响;其次,提出了基于序列二次规划算法的设计参数多目标优化方法,优化的目标包括高单位推力、低油耗和低红外辐射强度;最后,基于以上模型,利用序列二次规划算法对在设计点非加力情况下的涡扇发动机设计参数进行多目标优化。仿真结果表明:在设计点上,相较于不带直二元喷管的涡扇发动机,带直二元喷管的涡扇发动机具有更好的红外性能,并且通过算法优化后,带直二元喷管的涡扇发动机具有更好的性能参数。 相似文献
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N2O/HTPB固液火箭发动机喷管两相流计算 总被引:5,自引:5,他引:0
利用二维轴对称N-S方程对选用氧化亚氮/丁羟基燃料推进剂的固液混合火箭发动机的喷管两相流进行了计算.计算采用MacCormack时间推进预报校正二步格式,采用了Baldwin-Lomax代数湍流模型和两相平衡流模型.计算了三种氧燃比下4个不同喷管的喷管流场参数,并计算了喷管性能,通过比较两相流和气相流的计算结果,分析了不同氧燃比和喷管形状对喷管性能的影响,认为固液火箭发动机的性能主要受氧燃比的影响,为固液混合火箭发动机的设计提供了依据. 相似文献
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液体火箭发动机推力室的烧蚀冷却是一种新的冷却技术.本文介绍了烧蚀冷却的机理、变推力火箭发动机推力室中换热系数的计算、烧蚀速率和侵蚀速率的计算、室壁中温度的计算以及由于烧蚀冷却所引起的性能损失的计算等,可供设计烧蚀冷却推力室作参考. 相似文献
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Continuous Detonation Engine and Effects of Different Types of Nozzle on Its Propulsion Performance 总被引:4,自引:1,他引:3
The rotating propagation of a continuous detonation engine (CDE) with different types of nozzles is investigated in three-dimensional numerical simulation using a one-step chemical reaction model. Flux terms are solved by the so-called monotonicity-preserving weighted essentially non-oscillatory (MPWENO) scheme. The simulated flow field agrees well with the previous experimental results. Once the initial transient effects die down, the detonation wave maintains continuous oscil-latory propagation in the annular chamber as long as fuel is continuously injected. Using a numerical flow field, the propulsion performance of a CDE is computed for four types of nozzles, namely the constant-area nozzle, Laval nozzle, diverging nozzle and converging nozzle. The gross specific impulse of the CDE ranges 1 540-1 750 s and the mass flux per square meter ranges 313-330 kg/(m2&;#8226;s) for different nozzles. Among these four types of nozzles, Laval nozzle performs the best, and these parame-ters are 1 800 N, 1 750 s and 313 kg/(m2&;#8226;s). A nozzle can greatly improve the propulsion performance. 相似文献
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Thrust chamber injector is very critical component since that small differences in its design can result in dramatically different performance. This paper presents a preliminary investigation of a mono-propellant jet–swirl nozzle for thrust chamber using the existing commercial nozzle made of brass with little modification. The performance of two types of spray pattern, i.e. hollow cone and solid cone were investigated under cold flow test. Two solid cone injectors were studied at various purely axial orifice diameter with all other parameters remaining constant. This investigation reveals that decrease in the breakup length increased the spray angle. Higher injection pressure leads to shorter breakup length and higher value of discharge coefficient, except for hollow cone nozzle which shows mild fluctuating trend. Experimental data also tells that solid cone nozzle with diameter ratio 0.14 has the highest average discharge coefficient (0.60), followed by hollow cone nozzle (0.25). Solid cone nozzle with diameter ratio of 0.1 has an average discharge coefficient of 0.21. 相似文献
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固体火箭发动机碳基材料喷管热化学烧蚀特性 总被引:3,自引:3,他引:0
为了准确预示固体火箭发动机碳基材料喷管的烧蚀率,依据热化学烧蚀理论,建立了喷管传热烧蚀的二维轴对称气-固-热耦合计算模型,计算通过FLUENT壁面化学反应模型完成,无需事先假设烧蚀控制机制。针对70-lb BATES发动机喷管进行了烧蚀计算,研究了推进剂配方、氧化性组分、燃烧室压强对喷管烧蚀的影响。结果表明:烧蚀率计算值与试验测试值吻合较好;烧蚀率分布遵循喷管内壁热流密度分布规律,在喉部上游入口处达到峰值;烧蚀率随推进剂Al含量增加而降低,随燃烧室压强升高而近似正比例增大;H2O是决定烧蚀的主要氧化性组分。 相似文献
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《中国航空学报》2016,(3):630-638
Spray cooling has proved its superior heat transfer performance in removing high heat flux for ground applications.However,the dissipation of vapor–liquid mixture from the heat surface and the closed-loop circulation of the coolant are two challenges in reduced or zero gravity space environments.In this paper,an ejected spray cooling system for space closed-loop application was proposed and the negative pressure in the ejected condenser chamber was applied to sucking the two-phase mixture from the spray chamber.Its ground experimental setup was built and experimental investigations on the smooth circle heat surface with a diameter of 5 mm were conducted with distilled water as the coolant spraying from a nozzle of 0.51 mm orifice diameter at the inlet temperatures of 69.2 °C and 78.2 °C under the conditions of heat flux ranging from 69.76 W/cm~2 to 311.45 W/cm~2,volume flow through the spray nozzle varying from 11.22 L/h to 15.76 L/h.Work performance of the spray nozzle and heat transfer performance of the spray cooling system were analyzed;results show that this ejected spray cooling system has a good heat transfer performance and provides valid foundation for space closed-loop application in the near future. 相似文献
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