共查询到19条相似文献,搜索用时 40 毫秒
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针对火箭飞行工作中高空发动机燃料主管路系统防热罩存在热防护能力不足的问题,开展了高空羽流条件下的仿真计算和分析,依据温度计算值确定了防热罩紧固件在高温下抗拉伸强度低,在较高拧紧力矩条件下存在锌、镉脆断裂的薄弱环节,从而导致防热罩脱落。防热罩脱落后其内充填的隔热包覆材料被羽流吹落,燃料主汽蚀管连接法兰直接暴露在高温羽流环境中,高温导致法兰连接及密封失效从而产生燃料泄漏。针对防热罩热防护设计中存在的薄弱环节完成了设计改进,采用头锥形防热罩、高温合金材料的紧固件和多层耐高温隔热材料捆扎包覆等设计改进方案后,经过了高温、振动、地面发动机热试车和飞行试验验证,未出现前述故障。 相似文献
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25N双组元发动机热控研究 总被引:4,自引:0,他引:4
先前的推进系统25 N双组元发动机头部仅一个安装法兰盘,无支架,发动机长时间工作后法兰盘热反浸温度较高,不利于法兰盘上游电磁阀的工作性能。目前推进系统采用双法兰盘支架结构的新型25 N双组元发动机,由于新增支架的隔热,给热控带来了一定难度。在空间极端低温环境下,为使发动机温度满足点火前指标要求,须采取一定的热控措施。以25 N双组元发动机为研究对象,运用I-DEAS/TMG有限元热分析软件,建立了物理模型,研究了大小法兰盘在不同加热功率组合下发动机头部温度场的分布,并根据计算结果选择最佳加热功率组合。同时,根据经验配以适当的被动热控措施。通过飞行试验验证25 N发动机热控设计可靠性高,该热控设计方案可用于其他在研型号的推进系统。 相似文献
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铌铪合金表面硅化物涂层的高温失效行为分析 总被引:1,自引:0,他引:1
铌铪合金为轨姿控液体火箭发动机推力室身部主要结构材料,在高温有氧的工作环境中易发生氧化粉化,必须在合金表面涂覆高温抗氧化涂层。通过分析铌铪合金表面硅化物涂层的高温氧化、高温热震、瞬时高温烧蚀和热试车行为,阐述高温条件下的氧化失效行为。试验结果为:涂层1 800℃以下氧化条件下,表面形成致密的二氧化硅氧化膜,使得涂层的氧化寿命大于2 h;1 800℃以上的超高温氧化条件下,高温热冲击作用,涂层内部形成大量的烧蚀型网格结构,表面未形成二氧化硅氧化膜,氧化寿命小于10 s;热试车考核中,涂层满足推力室外壁面温度1 350℃以下的使用工况,抗氧化能力较好,随着氧化温度升高,涂层高温抗氧化能力迅速衰减。 相似文献
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铌铪合金具有较高的高温强度,是轨姿控液体火箭发动机推力室身部的主要结构材料,但在工作环境中易发生氧化“粉化”,必须在合金表面涂覆高温抗氧化涂层.本文主要研究了硅化物涂层对铌铪合金热防护行为,包括涂层的成型过程、高温抗氧化行为及高温抗热震行为等.试验结果为:涂层在1 700℃下的氧化寿命7 h,1 400~800℃的空冷热震循环次数4 700次,表面粗糙度30~60 μm.并对铌铪合金推力室身部涂层热试车情况进行了详细分析研究,对涂层在富氧高温燃气冲刷作用下的工作机理进行研究分析,总结了硅化物涂层的热防护机理,研究的新型硅化物涂层在高温条件下具有较好的性能. 相似文献
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对某一发动机地面热试车液路固有频率及低频振动进行了分析.建立了集中参数模型,对地面试车时和飞行时的液路一阶、二阶纵向固有频率进行了计算.得出的结果与实际试验数据吻合.在计算分析的基础上,对发动机低频振动机理进行了分析.计算与分析表明:蓄压器参加发动机地面热试车会造成试车系统液路纵向固有频率的改变,由于改变了的液路纵向固有频率没有经过大量的实际试车考验,因而不能够排除引起低频耦合振动、并且导致试车故障的可能性.建议对于蓄压器参加发动机地面热试车应该慎重考虑.如果确实需要蓄压器参加发动机地面热试车,从安全角度出发,应该进行进一步的理论分析、试验工作.本文的数学模型对低频振动特性分析以及安全分析具有借鉴价值. 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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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