共查询到17条相似文献,搜索用时 140 毫秒
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针对小推力过氧化氢/煤油推力室催化分解点火进行研究.作为过氧化氢/煤油双组元发动机的技术途径之一,还可扩展应用于催化分解点火火炬、补燃循环发动机中.推力室采用细颗粒催化剂床分解90%浓度过氧化氢(90%H2O2),分解的高温燃气使煤油雾化、蒸发和点火并且自维持燃烧.研究工作包括了催化剂床和气液喷注器的设计、单组元分解特性、双组元点火可靠性、工作效率及稳定性研究.试验中采用热容式燃烧室,催化剂床采用轮毂式分配板和多孔式床支板,并检验了不同结构的分解燃气与燃料喷射、混合情况.研究结果显示,催化分解点火可靠性高,工作稳定,燃烧效率在95%以上. 相似文献
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RBCC推进系统主火箭发动机气氧/煤油推力室研究 总被引:1,自引:0,他引:1
为满足RBCC推进系统主火箭发动机对气氧/煤油推力室的要求,对其进行了高燃烧室压力和温度、大范围变工况工作研究。气氧/煤油推力室喷注器采用中心区气液双组元内混式喷嘴和边区直流喷嘴结合结构,身部采用夹层冷却结构。通过对推力室气氧/煤油推进剂的点火及雾化混合技术、推力室喷注器及身部冷却设计技术、推力室的点火启动、稳态工作等关键技术的研究表明,推力室在室压3MPa、5MPa工况下可稳定燃烧。额定推力650N的气氧/煤油推力室方案可靠、点火工作正常,可以满足大范围变工况稳定工作要求。 相似文献
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我国新一代大推力液氧/煤油补燃发动机采用双推力室方案,发动机起动时存在推力室点火不同步情况.以500 t级液氧/煤油补燃发动机为研究对象,针对起动时推力室点火不同步问题,对发动机推力室燃料路的控制方案进行了研究.建立了描述补燃循环发动机起动过程的数学模型,搭建了双推力室发动机起动仿真平台.通过对推力室燃料路两种控制方案的对比分析:指出了从降低发动机系统对双推力室不同步点火的敏感程度考虑,采用2个燃料节流阀分别控制各分支燃料路的方案较优;推力室燃料路采用一个燃料节流阀的控制方案时,推力室冷却套流阻偏差不宜大于1 MPa. 相似文献
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推进剂的无毒化是航天发展的必然趋势,与肼类和硝基推进剂组合相比,过氧化氢/醇类具有低毒、廉价和可维护性好等优点,是未来最具竞争力的双组元绿色推进剂。过氧化氢/醇类最大缺点是不自燃,通过在醇类燃料中加入催化剂和添加剂解决了这一关键技术。利用自行设计的一套着火延滞期测定装置对含不同催化剂和添加剂的燃料配方与过氧化氢的自燃特性进行了测试、比较和评价。研究表明,过氧化氢/醇类这种新型双组元推进剂具有良好的自燃和点火特性。 相似文献
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现在高性能可长期贮存的火箭发动机燃料和推进剂的研究方向是开发能使用胶体燃料和胶体推进剂的燃烧系统。这种观点的产生是由于在液体燃料中加入胶状添加剂和(或)金属添加剂能显著提高液体推进剂的性能、密度和比冲。掌握胶体燃料单个液滴的蒸发和燃烧过程是预测未来推力室燃烧性能的第一个基本步骤。研究胶体燃料雾滴蒸发燃烧时应用了先进的计算机辅助摄像系统。测量了雾滴的燃烧速率,并精确观测研究了燃烧过程。对室压和氧化剂质量百分比浓度对雾滴燃烧速率的影响进行了量化研究.结果表明肢体燃料比液体燃料燃烧速率慢,而且点火更加困难。 相似文献
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提高轨控发动机的真空比冲可以有效减少卫星变轨推进剂的消耗量,从而延长卫星的在轨工作寿命或增加有效载荷质量。介绍了我国在研的卫星用第三代铼/铱材料490 N发动机设计方案、技术攻关和试验情况,对工程化应用存在的问题进行了分析,并提出了改进和优化方案。在第二代490 N发动机的设计基础上,第三代490 N发动机成功攻克了可靠传热稳定工作喷注器、高性能喷注器与燃烧室匹配以及新型高温抗氧化材料制备等关键技术,真空比冲提高了10 s,达到325 s。两台发动机均通过了25 000 s鉴定级高空模拟热试车寿命考核,性能指标达到国际先进水平。但是针对试车子样数较少和铼/铱燃烧室制备工艺困难的问题,仍需进一步开展铼基体和铱涂层的高温性能研究,并继续优化发动机设计。 相似文献
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Systems analysis of a Mach 5 class hypersonic aircraft is performed. The aircraft can fly across the Pacific Ocean in 2 h. A multidisciplinary optimization program for aerodynamics, structure, propulsion, and trajectory is used in the analysis. The result of each element model is improved using higher accuracy analysis tools. The aerodynamic performance of the hypersonic aircraft is examined through hypersonic wind tunnel tests. A thermal management system based on the data of the wind tunnel tests is proposed. A pre-cooled turbojet engine is adopted as the propulsion system for the hypersonic aircraft. The engine can be operated continuously from take-off to Mach 5. This engine uses a pre-cooling cycle using cryogenic liquid hydrogen. The high temperature inlet air of hypersonic flight would be cooled by the same liquid hydrogen used as fuel. The engine is tested under sea level static conditions. The engine is installed on a flight test vehicle. Both liquid hydrogen fuel and gaseous hydrogen fuel are supplied to the engine from a tank and cylinders installed within the vehicle. The designed operation of major components of the engine is confirmed. A large amount of liquid hydrogen is supplied to the pre-cooler in order to make its performance sufficient for Mach 5 flight. Thus, fuel rich combustion is adopted at the afterburner. The experiments are carried out under the conditions that the engine is mounted upon an experimental airframe with both set up either horizontally or vertically. As a result, the operating procedure of the pre-cooled turbojet engine is demonstrated. 相似文献
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Test results of the air turbo ramjet for a future space plane 总被引:1,自引:0,他引:1
The Institute of Space and Astronautical Science (ISAS) has been engaged in the development study on the Air Turbo Ramjet (ATR) engine since 1986 in cooperation with the Ishikawajima Harima Heavy Industries Co. Ltd (IHI). The ATR is one of the most preferable candidates for the propulsion system of a future space plane. Our ATR engine is a combined cycle air breathing propulsion system which consists of the turbojet and the fan boosted ramjet using the liquid hydrogen as a fuel. This engine system was named “ATREX” after employing the expander cycle. The ATREX is energized by thermal energy extracted regeneratively in both the pre-cooler installed in the air intake and the heat exchanger in combustion chamber. The ATREX works in the flight condition from sea level static up to Mach 6 at 35 km altitude. The ATREX employs the tip turbine configuration for compactness of turbo machinery. We are assessing the feasibility of the ATREX system by the sea level static tests using the 1/4-scale model (ATREX-500) with a fan inlet diameter of 300 mm and overall length of 2120 mm. In 1990, the ATREX-500 engine was tested in a sea level static condition to verify the performance characteristics of the turbo machinery and the combustor. In September of 1991, the heat exchanger was installed in the combustion chamber and tested independently from the turbo system. In November of 1991, the heat exchanger was coupled with the turbo system and tested to verify the overall system of the ATREX. In this paper are presented the test results of the ATREX-500 engine tested in the sea level static condition. 相似文献
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燃烧室内壁是液氧煤油发动机推力室中的重要部件之一。其工作时承受高温、高压燃气;内、外壁均为曲线拟合母线的回转体;直径大、长度长、且壁薄;内冷却环带槽结构特殊。在整个研制过程中,主要就加工方法、工艺流程、零件的装夹、定位基准的确定、程序设计、工艺装备设计、切削刀具、检测量具、切削参数等方面做了研究。按研究方案已加工出多件产品,经整机系统试车,效果良好,达到了预期目的和效果。 相似文献