共查询到20条相似文献,搜索用时 140 毫秒
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首先完成了一种典型DMSJ发动机流道型面和燃烧组织设计,该发动机在M_∞=4.0和6.0时的比冲分别为1 029.6 s和899.9 s。以此DMSJ发动机流道为基础,在隔离段一侧布置火箭发动机,形成RBCC发动机流道。数值模拟研究表明,低马赫数时,火箭台阶及下游流道型面变化对发动机性能影响有限;保持DMSJ发动机燃料喷注方案不变,RBCC发动机在M_∞=4.0时,冲压模态比冲可达到1 052.8 s。高马赫数时,由于燃烧组织位置靠前,必须对DMSJ发动机原有的燃料喷注方案进行调整,才能确保RBCC发动机达到与前者相当的比冲水平,经过调整本文RBCC发动机M_∞=6.0时冲压比冲达到了887.8 s。因此,基于目前较成熟的DMSJ发动机进行高马赫数RBCC发动机设计,是一条快速可行的技术途径。 相似文献
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火箭引射模态下一次火箭流量优化方法研究 总被引:4,自引:0,他引:4
对火箭基组合循环(RBCC)发动机火箭引射模态下一次火箭流量优化方法开展了研究,并对飞行条件下一次火箭流量的变化规律进行了分析。提出了采用有效比冲作为优化目标的一次火箭流量单目标优化模型;建立了求解与一次火箭流量相匹配的二次燃料流量的比例控制方法;在考虑发动机性能优化与弹道分析耦合作用的基础上,采用试验设计和遗传算法,建立了火箭引射模态下一次火箭流量优化方法。针对空中载机发射的RBCC发动机,开展了火箭引射模态下一次火箭流量优化,并根据弹道分析结果,给出了飞行条件下一次火箭流量变化规律。结果表明,为了克服飞行过程中声障阻力,一次火箭流量在Ma=1.0附近达到最大,此时对发动机提出较高的推力设计要求;在Ma=1.5附近,来流空气的冲压作用占主导地位,一次火箭流量出现较大程度的节流,此时对发动机提出较高的比冲设计要求;超过Ma=1.5后,一次火箭以较小的流量状态维持稳定工作;火箭引射模态下一次火箭流量调节比达到了5.0。 相似文献
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介绍了采用引射火箭模式的RBCC发动机工作原理,并在对其概念设计模型进行简化的基础上,进行了RBCC发动机系统性能分析,评估了RBCC发动机系统主要设计参数(发动机系统出口截面直径和燃料化学反应后的总温)的变化对其性能(推力、推力系数和比冲)的影响,认为:1)燃料经过加热后,推力和推力增益都上升了69.97%,比冲增加了180.18%;2)随着二次燃烧过程中燃烧室温度的上升,发动机的推力、推力增益和比冲得到了很大的提高,火箭发动机的性能得到了很好的改善;3)随着RBCC发动机系统出口截面直径的增加,尾气对发动机的反推力、推力增益以及比冲急剧下降,不利于其性能的改善。 相似文献
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氢氧火箭发动机因具有很高的性能,在国内外运载火箭中得到了广泛应用。为了更好地开展深空探测,必须研制大推力氢氧火箭发动机。本文综合分析了国内外氢氧发动机的发展历史和现状,简要介绍了我国220 t补燃循环氢氧发动机方案和关键技术研制进展,该技术方案先进。通过开展试验件冷态试验和缩尺组件热试验等,研究了核心部组件的关键技术和方案选型。开展全尺寸预燃室热试验等分系统热试验,研究组件级技术,初步突破了部分关键技术。建议加快开展大推力氢氧发动机工程研制,促进我国航天推进技术发展。 相似文献
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The feasibility study was conducted to use the cryogenic propulsion system for the third stage of the future H-1 vehicle. While the third-stage mass fraction is less than the current solid propellant third stage, the 50% higher Isp results in a significantly higher payload. Two basic configurations of the new propulsion system were proposed: one pressure-fed system and two pump-fed systems. The first is a pressure-fed system providing a 700 kg thrust at an Isp of 441 sec with restart capability. The second is a pump-fed system, operating on an expander cycle principle. A midget turbopump with a 90 000 rpm shaft speed feeds the thrust chamber which delivers 1 ton of thrust at an Isp of 471 sec. The third proposed system is also a pump-fed design using a unique expander bleed cycle, and delivers a 1 ton thrust at an Isp of 470 sec with a turbopump speed of 80 000 rpm. The results of engine testing predict the performance feasibility of respective propulsion system designs. 相似文献
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为研究烃类推进剂航天动力技术在中国的后续发展和未来应用方向,对比分析煤油、甲烷和丙烷等典型烃类推进剂的物理化学性质和应用特性,简要介绍烃类推进剂航天动力在一次性运载火箭、可重复使用运载器、高性能上面级推进、无毒空间推进和吸气式推进领域的发展动态及应用状况。当前国内外航天动力系统的发展和应用情况表明,以液氧煤油发动机和液氧甲烷发动机为代表的烃类推进剂航天动力将引领未来高性能低成本航天推进系统的发展趋势,依照中国液氧/烃火箭发动机的研制进展和技术水平,以其为核心的新型动力体系在中国未来的天地往返、载人登月和深空探测等多任务适应性方面具有良好应用前景。 相似文献
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The pressure-fed second stage propulsion system for N-launch vehicle provides 53,348 N (5440 kg) thrust for about 250 sec at an Isp of 290.2 sec. Aluminum tanks, integral with vehicle structure, carry a minimum of 4.7 ton propellant combination of N2O4 and Aerozine 50. The gimbaled engine consists of a regenerative cooled chamber, ablative nozzle spacer, and a radiation cooled nozzle extension with an exit area ratio of 26. Utmost utilization of domestically available technology and facilities underlay the design concept. Development of the propulsion system took 5 years with the first flight occurring in 1975. Five consecutive flight successes up to 1979 have demonstrated the reliability and performance of the system.Improved N vehicle, designated as N-II, will succeed the N vehicle. New second stage propulsion system for N-II delivers 43,816 N (4468 kg) thrust at an Isp of 314.1 sec and has restart-capability. 相似文献
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载人小行星探测的任务特点与实施途径探讨 总被引:2,自引:1,他引:1
介绍了载人小行星探测的发展现状,对目前美国基于"猎户座"飞船的载人小行星探测的概要方案进行了描述,包括探测器系统组成、运载火箭和飞行方案等内容。从速度增量、目标星引力等方面,分析了载人小行星探测的任务特点,并与载人火星探测、载人月球探测以及无人小行星探测的任务特点进行了比较。给出了载人小行星探测的实施途径建议,包括目标星选择、载人飞船系统设计等。讨论了其所涉及的推进、星际飞行安全保障、小行星表面行走等关键技术。研究结果可为我国开展载人深空探测提供参考。 相似文献
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Power-limited systems with variable Isp, which have been studied theoretically since the beginning of astronautics, are getting closer to practical applications thanks to recent technological advances in the field of magnetosplasma rockets, such as Ad-Astra’s VASIMR concept. This type of propulsion system is considered for high-speed interplanetary transfers, such as Mars missions, with demanding payload fractions that would be compatible with manned missions. This paper explores the problem of the optimization of a power-limited propulsion system through simple performance models, and investigates the trade-off between the technological requirements, the transfer time and the payload fraction1. Following previous works existing in literature, we model the technological characteristics of the vehicle through a small number of parameters, the most important of which being the specific weight (or mass-to-power ratio) of the power generation system. Also, we use in our models the classical “trajectory characteristic” parameter (defined as the integral over time of the squared thrust acceleration) which represents – under certain hypotheses – the propulsion requirements for an orbital or interplanetary transfer with a given time and a given thrust strategy. In this paper, we first give a review of existing methods in literature, then we present the equations of a new class of optimal design which maximizes the payload fraction, for a given transfer time and given technological characteristics. This class of optimal design is described through very simple equations that make possible to study more straightforwardly than existing calculations the links between the main mission requirements (transfer time and payload fraction) and the main technological requirements (specific weight of the power generation and structure mass ratio of the whole vehicle, excluding the power generation system). One important result obtained from these equations is a simple expression which estimates the theoretical upper limit of the power source’s specific weight as a function of transfer time and the payload mass ratio. In the last part of this paper, we apply this simple performance model to discuss the feasibility of a fast Earth-to-Mars transfer using a power-limited system. 相似文献
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深空探测作为我国航天领域未来的重要任务之一,需要性能更高的推进系统提供动力。核热推进系统具有高比冲、大推力、长运行寿命、可重复启动等优点,可为未来深空探测任务提供可靠的动力支撑。经过了60多年的发展,核热推进固态堆芯燃料元件被研制出了多种类型,如六棱柱石墨基燃料元件、扭曲条带燃料元件、六棱柱金属陶瓷燃料元件、球形包覆颗粒燃料元件、MITEE型燃料元件、SLHC型燃料元件、Grooved Ring型燃料元件等。总结归纳了核热推进固态堆芯燃料元件的发展状况,提出了发展核热推进固态堆芯燃料元件的关键技术,可为我国核热推进系统燃料元件的研制提供借鉴。 相似文献