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针对宽范围工作的火箭基组合循环(RBCC)发动机推力、比冲和燃烧效率等性能难以通过关键参数的实时测量直接计算的问题,提出利用可调谐二极管激光器吸收光谱(TDLAS)技术实现发动机燃气温度、H2O组分浓度和速度多参数同时实时在线非接触式激光光谱测量方法,搭建用于RBCC地面试验性能分析的时分复用-扫描波长TDLAS系统。通过试验获得的(7444.352+7444.371)/7185.597cm -1 谱线附近吸收光谱,从而得到燃气温度、H2O组分浓度和速度参数,并结合数值模拟方法确定的流场参数纵向分布,实现基于关键参数在线测量的发动机性能直接计算。该方法有助于评判燃烧组织和结构改变对发动机性能的影响。 相似文献
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《中国航空学报》2020,33(5):1454-1467
In this paper, an Uncertainty-based Multi-disciplinary Design Optimization (UMDO) method combining with fuzzy theory and Multi-Discipline Feasible (MDF) method is developed for the conceptual design of a Hybrid Rocket Motor (HRM) powered Launch Vehicle (LV). In the method proposed, membership functions are used to represent the uncertain factors, the fuzzy statistical experiment is introduced to analyze the propagation of uncertainties, and means, standard deviations and credibility measures are used to delineate uncertain responses. A geometric programming problem is solved to verify the feasibility of the Fuzzy-based Multi-Discipline Feasible (F-MDF) method. A multi-disciplinary analysis of a three-stage HRM powered LV involving the disciplines of propulsion, structure, aerodynamics and trajectory is implemented, and the mathematical models corresponding to the F-MDF method and the MDF method are established. A two-phase optimization method is proposed for multi-disciplinary design optimization of the LV, including the orbital capacity optimization phase based on the Ziolkowski formula, and the scheme trajectory verification phase based on the 3-degree-of-freedom point trajectory simulation. The correlation coefficients and the quadratic Response Surface Method (RSM) based on Latin Hypercube Sampling (LHS) are adopted for sensitive analysis of uncertain factors, and the Multi-Island Genetic Algorithm (MIGA) is adopted as the optimization algorithm. The results show that the F-MDF method is applicable in LV conceptual design, and the design with the F-MDF method is more reliable and robust than that with the MDF method. 相似文献
195.
王广民 《中国空间科学技术》1990,10(3):42-48
文章提供利用运载火箭特征点(关机点,入轨点等)的有关加速度计的积分遥测量计算该点的弹道参数和轨道要素的计算方法。 相似文献
196.
Aerospike nozzle contour design and its performance validation 总被引:1,自引:0,他引:1
A simplified design and optimization method of aerospike nozzle contour and the results of tests and numerical simulation of aerospike nozzles are presented. The primary nozzle contour is approximated by two circular arcs and a parabola; the plug contour is approximated by a parabola and a third-order polynomial. The maximum total impulse from sea level to design altitude is adopted as objective to optimize the aerospike nozzle contour. Experimental studies were performed on a 6-cell tile-shaped aerospike nozzle, a 1-cell linear aerospike nozzle and a 3-cell aerospike nozzle with round-to-rectangle (RTR) primary nozzles designed by method proposed in present paper. Three aerospike nozzles achieved good altitude compensation capacities in the tests and still had better performance at off-design altitudes compared with that of the bell-shaped nozzle. In cold-flow tests, 6-cell tile-shaped aerospike nozzle and 1-cell linear aerospike nozzle obtained high thrust efficiency at design altitude. Employing gas H2/gas O2 (GH2/GO2) as propellants, hot-firing tests were carried out on a 3-cell aerospike nozzle engine with RTR primary nozzles. The performance was obtained under two nozzle pressure ratios (NPR) lower than design altitude. Efficiency reached 92.0–93.5% and 95.0–96.0%, respectively. Pressure distribution along plug ramp was measured and the effects of variation in the amount of base bleed on performance were also examined in the tests. 相似文献
197.
针对模态过渡过程中的瞬变特性,开展了引射/亚燃模态过渡工作过程中的RBCC进气道/燃烧室一体化数值模拟,计算比较了四种不同的模态过渡方案,分析了模态过渡工作过程中流动、燃烧模式和进气道状态的瞬变过程,研究结果表明:(1)推力产生较大波动的主要原因是燃烧室内燃烧/流动参数匹配性差和在流动方向上燃烧放热间断引起的;(2)主火箭保留的燃气在模态过渡过程中起到了火焰稳定和自持燃烧的作用,保留合理流量的燃气不仅可以缩短模态过渡时间,而且可以提高发动机的比冲;(3)提出了通过调节燃料喷注策略和主火箭节流方式实现模态平稳过渡的方案,并对该方案进行了数值验证,可望为进一步的实验研究提供了燃烧控制方法。 相似文献
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