Detonation engine fed by acetylene–oxygen mixture

The advantages of a constant volume combustion cycle as compared to constant pressure combustion in terms of thermodynamic efficiency has focused the search for advanced propulsion on detonation engines. Detonation of acetylene mixed with oxygen in various proportions is studied using mathematical modeling. Simplified kinetics of acetylene burning includes 11 reactions with 9 components. Deflagration to detonation transition (DDT) is obtained in a cylindrical tube with a section of obstacles modeling a Shchelkin spiral; the DDT takes place in this section for a wide range of initial mixture compositions. A modified ka-omega turbulence model is used to simulate flame acceleration in the Shchelkin spiral section of the system. The results of numerical simulations were compared with experiments, which had been performed in the same size detonation chamber and turbulent spiral ring section, and with theoretical data on the Chapman–Jouguet detonation parameters.     

面向平行系统的飞行器航迹仿真模型架构设计

面向平行系统运行需求,对飞行器航迹计算的积分、抽象计算、数学模型等进行建模和架构设计,实现了航迹计算过程中数据与算法的分离,并采用继承和指针注册机制完成与航迹计算相关的数据统一化分层管理,通过对飞行器状态转移模型的设计实现了飞行器不同状态或阶段之间航迹的顺利衔接,能够实现多状态并行转移和执行,以抵消平行系统不确定性和随机性的影响,为平行系统运行提供灵活、高效的仿真模型。     

Effects of a hot jet on detonation initiation and propagation in supersonic combustible mixtures

The effects of a hot jet on detonation initiation and propagation in supersonic combustible mixtures has been studied with two-dimensional numerical simulations with the open-source program AMROC that uses a block-structured adaptive mesh refinement method. Results indicate that the hot jet could ignite the detonation effectively in supersonic combustible mixtures like a pneumatic ramp. After the realization of the detonation initiation, the hot jet can still play an important role on the detonation propagation during its continuous ejection. For a hot jet with certain diameter, it can result in an overdriven detonation with almost constant overdrive degree. After the shutdown of the hot jet, the stable CJ detonation combustion was realized finally in the supersonic combustible mixtures. With the re-ejection of the hot jet, the failed detonation could be reinitiated quickly. Through the control of the re-ejection of the hot jet, it plays a key role not only in the initiation process, but also in the subsequent continuous detonation combustion period.     

Some technical aspects of gaseous detonation

This paper briefly describes two attempts to utilize detonative combustion processes to MHD conversion of thermal energy of fuel to electrical energy and bonding of atmospheric nitrogen. For this purpose a continuous impulse detonation chamber with a frequency up to 200 cps was constructed. Using methane-oxygen-nitrogen mixtures the chamber was maintained in stable operation for several hundred hours. Oil was also employed as fuel.Estimates based on experimental data showed that up to 2% of chemical energy of the fuel may be converted into electrical energy. The use of an accelerating nozzle may improve this result.The concentration of nitrogen oxide in combustion products of the detonation wave was higher by 14% than that expected under usual combustion conditions.The advantages of this type of apparatus are: absence of compressors for fuel and oxidant, impulse current generation, low temperatures of chamber walls, and operation over a large range of operating conditions.Problems associated with the effect of the magnetic field on the propagation of the detonation wave are discussed and the possibility of applying the Zeldovich theory to the case of MHD interaction is described. It is shown that the detonation velocity may either increase or decrease depending on the relative orientation of the direction of magnetic field with respect to the detonation wave.     

Super-ion-beams and advanced target concepts for thermonuclear microbomb rocket propulsion

Expected advances in the generation of ultraintense ion beams with currents above the Alfvén limit will make possible the ignition of neutron-poor advanced thermonuclear reactions suitable for thermonuclear microbomb propulsion. The superbeams can be produced by magnetically insulated multistage pulse accelerators. The high thermonuclear yields as they are desirable for an efficient propulsion system can be obtained by target staging and autocatalytic detonation. This will make possible the fast economical transportation of large payloads within the solar system.     

爆震波点火器在氢氧塞式喷管的工程应用

建立了氢氧爆震波点火器试验系统,并根据试验塞式喷管发动机工作状态要求设计了爆震波点火器。在高空条件下(0.005￣0.002MPa),爆震波点火器供气压力0.3MPa、混合比3左右,对爆震波点火器的点火性能进行了试验,成功实现了高空条件下爆震波点火火炬。在同样高空条件下对爆震波点火器点燃单元塞式喷管试验发动机成功进行了点火试验。试验结果表明,氢氧爆震波点火器能以较低的供气压力实现可靠点火。爆震波点火器在气氢气氧单元塞式喷管试验发动机点火的成功应用,为下一阶段应用于多管塞式喷管发动机的实际点火试验提供了技术基础。     

气相氛围中悬浮粉末燃料爆震燃烧研究进展

针对在氧化性气相氛围以及在燃料/氧化剂混合气相氛围中粉末燃料爆震燃烧波的传播特性,总结了气相氛围中悬浮粉末燃料爆震燃烧的实验和数值模拟研究进展,归纳了影响爆震波速度、稳定性、传播模式、细观结构和胞格尺寸的主要因素。同时,还介绍了粉末燃料应用于爆震发动机或燃烧室的案例;结合粉末爆震自身特点对实验装置和燃烧诊断测试手段和数值模拟方法进行总结分析;最后针对下一步需要开展的研究工作进行展望 。     

连续爆轰发动机的研究进展

连续爆轰发动机是一种基于爆轰波将推进剂的化学能转化成热能的新概念发动机,近年来受到世界各主要国家的高度关注。现已成功获得多种燃料长时间稳定的连续爆轰,较深入地认识了连续爆轰流场结构,初步测得推力和比冲,验证了连续爆轰发动机的性能优势并在火箭模态、冲压模态以及涡轮模态下都实现了稳定连续爆轰。对连续爆轰发动机的工作原理,以及近年来世界各主要国家在连续爆轰发动机的基础研究和应用研究方面取得的代表性成果进行了综述,并给出尚待解决的问题,为其进一步工程化应用提供参考。     

微纳卫星混合推进技术研究综述

首先介绍了微纳卫星混合推进系统的特点与优势,总结了星载混合推进技术的发展现状,并结合微纳卫星动力系统要求,指出了星载混合推进技术的发展趋势及存在的问题.然后在此基础上,从卫星工程应用的角度,重点综述了星载混合推进系统需攻克的主要关键技术,包括先进一体化制造、可靠低功耗多次点火启动、比冲提升、高效燃烧等,并分析了相应关键...     

小推力推进系统起动过程的分析

本文对小推力推进系统各部件建立了数学模型，并对此系统进行了数值计算。计算结果表明，在燃烧时滞较大时，该系统响应较慢，发动机参数的超调量较大，达到稳态所需的时间较长；轨控发动机与姿控发动机共用同一个供应系统时，姿控发动机受燃烧时滞的影响更大。减小燃烧时滞有利于提高发动机在起动过程的响应能力和稳定性。在起动阶段，高室压推进系统比低室压推进系统响应快，高室压轨控发动机的参数能较快地稳定下来，但其超调量较大；高室压姿控发动机虽然响应快，但其超调量大，达到稳态所需的时间长于低室压姿控发动机。本文所得结论为提高小推力推进系统在起动过程的响应能力提供了参考。     

光敏推进剂激光烧蚀过程建模与仿真

相较于传统大卫星,微小卫星具有结构紧凑、质量轻便和成本低廉的特点。然而,受功率和质量负载的限制,微小卫星一般不装备推进系统,其航线也局限于近地轨道。为扩展微小卫星的功能,满足日益复杂的任务需求,需给其配备合适的微推进系统。固体推进系统具有结构简单、寿命长、可靠性高的优点,但无法重复启动。为得到可重复启动的固体微推进系统,设计了一种非自持燃烧的光敏推进剂,采用激光控制其燃烧。在背压为大气压的环境下,利用高速摄像机拍摄燃烧过程并记录燃速。之后,对光敏推进剂的激光烧蚀过程进行建模。分析结果表明:激光可控制光敏推进剂的燃烧,燃速与激光强度成线性关系;该光敏推进剂的最小激光点火强度为0.28 W/mm~2;燃速计算值与实测值的误差在10%以内,证明该数学模型具备工程应用价值。     

航天器动力管路热控设计与试验研究

针对某航天器动力系统管路布局分散造成系统温差大、控温难的问题,结合动力管路温度指标要求和边界环境条件,采用以被动热控措施为主、辅以电加热主动热控措施的设计方案。分析确立动力管路的热环境,建立换热模型;通过仿真分析和整器热平衡试验,选取不同工况,验证了动力系统氧化剂管路和燃烧剂管路温度均维持在8～20℃范围内的热控设计结果。该方案对各类航天器的动力管路热控设计和分析有一定的指导和借鉴作用。     

混合比影响爆震波点火器工作过程之实验研究

爆震波点火器用于工程，其设计存在一个最佳结合点，使得在合适的管路中，爆震波传播速度、转捩距离、爆震波能量等能够符合点火器目标需求。为了研制适用于工程的爆震波点火器，在氢氧爆震波点火器基本特性试验的基础上，对初始混合气体的混合比等与爆震波特性的关系进行了研究。对实验结果进行分析认为。混合比对爆燃爆震转捩（DDT）距离影响较大，混合比大于3时，其转捩距离小于500mm。混合比增加时，爆震波传播速度会减小，但稳定的爆震波相对于波的混气的马赫数并小减小，维持在4．8左右。在初始混气压力不变情况下，质量流量可以提高爆震波能量，增强爆震波的点火能力。研究结论时爆震波点火器在工程中实际应用及以后的研究方向具有指导性作出。     

液体火箭凝胶推进剂燃烧特性研究进展

液体火箭凝胶推进剂燃烧特性是凝胶推进技术发展的关键问题之一。综述了液体火箭凝胶推进剂燃烧特性研究状况与进展,详细阐述了凝胶液滴燃烧机理,总结分析了凝胶推进剂燃烧特性三种试验研究方案和四种理论研究模型的特点,指出了研究凝胶推进剂燃烧特性的重要性并对进一步的研究工作提出建议。     

Benefits from incorporation of combined cycle propulsion

The X-33 program was initiated to develop a testbed for integrated RLV technologies that pave the way for a full scale development of a launch vehicle (Venture Star). Within the Nasa Future X Trailblazer program there is an Upgrade X-33 that focuses on materials and upgrades. The authors propose that the most significant gains can be realized by changing the propulsion cycle, not materials. The cycles examined are rocket cycles, with the combustion in the rocket motor. Specifically, these rocket cycles are: turbopump, topping, expander, air augmented, air augmented ram, LACE and deeply cooled. The vehicle size, volume, structural weight remain constant. The system and propellant tank weights vary with the propulsion system cycle. A reduction in dry weight, made possible by a reduced propellant tank volume, was converted into payload weight provided sufficient volume was made available by the propellant reduction. This analysis was extended to Venture Star for selected engine cycles. The results show that the X-33 test bed could carry a significant payload to LEO (10,000 Ib) and be a valuable test bed in developing a frequent flight to LEO capability. From X-33 published information the maximum speed is about 15,000 ft/sec. With a LACE rocket propulsion system Venture Star vehicle could be sized to a smaller vehicle with greater payload than the Venture Star baseline. Vehicle layout and characteristics were obtained from: http:// www.venturestar.com.     

弹丸速度对超爆轰冲压加速器性能影响

为研究超爆轰模态冲压加速器的推进性能,采用混合的Roe/HLL（Harten, Lax, Van Leer）格式,结合自适应网格加密技术(AMR )与沉浸边界法(IBM ),数值模拟了弹丸速度高于预混可燃气体C-J爆速的冲压加速器流场,揭示了弹丸速度对流场结构与推力的影响。结果表明当弹丸速度在一定范围时,斜爆轰波可驻定在弹丸肩部或头部,在弹丸尾部形成高压区加速弹丸,并且,斜爆轰波驻定在弹丸头部推力更高,稳定工作的速度范围 更宽 。     

液氧／煤油补燃循环发动机起动过程研究

液体火箭发动机起动过程是发动机研制过程中的难点和关键技术之一。针对某液氧／煤油补燃循环发动机,进行了起动过程研究。建立了发动机各组件的动态数学模型,并进行了适当简化。计算得到了起动过程发动机性能参数随时间变化的仿真曲线。计算结果与试车数据基本相符,初步验证了所建立的仿真模型及采用的仿真方法的正确性。还分析了部分干扰因素对发动机起动过程的影响。     

铝粉燃料与水反应的研究进展

Al/H2O反应具有高能量密度,可用于水下推进和太空发动机中。对水反应Al粉燃料的制备和Al/H2O燃烧反应的研究现状进行了综述。通过分析认为,Al的超细化和对其包覆处理能提高Al粉的抗氧化能力与水反应活性,可改善Al粉贮存性能及水反应的燃烧性能。阐述了Al粉点火、水反应燃烧机理及其影响因素,对Al/H2O反应进一步研究的重点及其应用前景进行了展望。     

膏体冲压发动机构型对燃烧效率的影响研究

膏体冲压发动机是在膏体发动机和冲压发动机基础上提出的一种新型组合动力装置.探讨了膏体冲压发动机的多种构型.并对其反应流场进行了数值模拟和燃烧效率分析.结果表明,采用将膏体富燃料直接喷射到补燃室的方式燃烧效率较高;随着进气道轴向距离的增加,补燃室头部的燃烧效率有所降低;随着进气道夹角的增加,补燃室燃烧效率略有增加.     

变轨发动机不等量截尾试验可靠度评估

对于可靠度要求极高和任务时间以小时计的变轨发动机来说，目前广泛采用以铌合金为材料并喷涂以抗高温氧化涂层方案的推力室，这一类发动机在方案试验阶段结束后的研制试验中往往只出现截尾试验的结果。给出了不等量截尾试验结果的可靠度评估方案。