航天推进技术研究院主办
[1]赵 婷,刘昌国,吴凌峰,等.双组元离心式喷注器10 N发动机偏工况试验[J].火箭推进,2020,46(01):52-58.
ZHAO Ting,LIU Changguo,WU Lingfeng,et al.Experimental investigations on non-rated condition performance of a 10 N thruster with bi-centrifugal swirl injector[J].Journal of Rocket Propulsion,2020,46(01):52-58.
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ZHAO Ting,LIU Changguo,WU Lingfeng,et al.Experimental investigations on non-rated condition performance of a 10 N thruster with bi-centrifugal swirl injector[J].Journal of Rocket Propulsion,2020,46(01):52-58.
双组元离心式喷注器10 N发动机偏工况试验
《火箭推进》[ISSN:1672-9374/CN:CN 61-1436/V]
卷:
46
期数:
2020年01期
页码:
52-58
栏目:
研究与设计
出版日期:
2020-02-29
- Title:
- Experimental investigations on non-rated condition performance of a 10 N thruster with bi-centrifugal swirl injector
- 分类号:
- V434.3文献标识码:A 文章编号:1672-9374(2020)01-0052-07
- 摘要:
- 根据国内外同类发动机研制经验,双组元10 N发动机在入口压力为0.8～2.2 MPa范围内,入口压力偏差会使发动机真空比冲、燃气温度等性能产生较大变化。为了获得双组元离心式喷注器10 N发动机在落压推进系统要求的入口压力范围内性能,通过采用小流量喷雾试验台和42 km 高模试验台,对偏工况条件下的冷态性能及热试性能进行试验研究。试验结果表明:该发动机额定入口压力1.58 MPa时真空比冲为2 881 N·s/kg; 当入口压力在0.6～2.5 MPa变化时,对应真空推力从4.7 N增加到14 N,落压比为3; 入口压力0.6 MPa时真空比冲为2 600 N·s/kg,入口压力2.5 MPa 时真空比冲为2 956 N·s/kg; 入口压力在0.6～2.5 MPa试验范围内,发动机燃烧室壁温均低于材料许用温度,表明发动机热设计优良,可满足双组元落压推进系统对姿控发动机的性能需求。
- Abstract:
- According to the development experience of 10 N bi-propellant thrusters at home and abroad, the range of inlet pressure is from 0.8 MPa to 2.2 MPa.The deviation of inlet pressure will cause great changes in thruster performance such as vacuum specific impulse and gas temperature.To investigate the performance of a 10 N bi-propellant thruster at different supply pressures from 0.8 MPa to 2.2 MPa,experiments were conducted on the discharge device with two-dimensional particle dynamic analyzer,as well as with the 42 km-high-altitude simulation test facility.It has successfully undergone all qualification tests recommended for bipropellant blowdown propulsion system.The results show that the 10 N bi-propellant thruster can fit the supply pressures from 0.6 MPa to 2.5 MPa,the nominal specific impulse of the thruster is greater than 2 881 N·s/kg when the inlet pressure is 1.58 MPa.When the inlet pressure changes from 0.6 MPa to 2.5 MPa, the corresponding vacuum thrust increases from 4.7N to 14 N and the pressure drop-down ratio is 3.As inlet pressure increases, the specific impulse also increases, when the inlet pressure is 0.6 MPa, the vacuum specific impulse is 2 600 N·s/kg; and when the inlet pressure is 2.5 MPa, the vacuum specific impulse is 2 956 N·s/kg.Within the range, the combustion chamber temperature of the thruster is lower than the allowable temperature of the material, reflecting the excellent design of the thruster, which can meet the performance demands of bipropellant blowdown propulsion system for attitude control.
参考文献/References:
[1] SCHULTE G, GOTZIG U,HORCH A,et al.Further improvements and qualification status of astrium’s 10 N bipropellant thruster family[R].AIAA 2003-4776.
[2] BALAN C G.Development of a 10 N liquid bipropellant thruster for geostationary spacecraft programme of Indian Space Research Organisation[R].AIAA 2009-5529.
[3] MARQUARDT K.Mars flyer rocket propulsion risk assessment[R].NASA CR 2001-210710.
[4] GORDON A.TRW pintle engine heritage and performance characteristics[R].AIAA 2000-3871.
[5] RICHARD J.Results from tests on a 10 N thruster using low temperature(-40℃)propellants [R].AIAA 2001-3254.
[6] WOOD R S.Development of a low cost 22 N bipropellant thruster[R].AIAA 1990-2056.
[7] WILLIS W D.The SDO propulsion subsystem[R].AIAA 2012-4329.
[8] 杨俊, 何永英, 连仁志, 等.双组元落压推进系统应用现状及关键技术[J].火箭推进, 2016, 42(4):21-25.YANG J, HE Y Y, LIAN R Z, et al.Application status and key technology of bipropellant blowdown propulsion systems[J].Journal of Rocket Propulsion, 2016, 42(4):21-25.
[9] 李春红.ARC公司22N双组元姿控发动机大落压工况性能验证[J].火箭推进, 2002, 28(3):52-59.LI C H.Performance verification of ARC 22N two-component attitude control engine under high drop pressure [J].Journal of Rocket Propulsion, 2002, 28(3):52-59.
[10] 毛晓芳, 汪凤山, 杨晓红, 等.小推力双组元姿控发动机性能研究[J].推进技术, 2012, 33(6):987-990.
[11] 刘志泉, 叶超, 林庆国.微小型双组元姿控发动机技术研究[J].火箭推进, 2014, 40(4):1-6.LIU Z Q, YE C, LIN Q G.Study on miniature liquid bipropellant attitude control rocket engine[J].Journal of Rocket Propulsion, 2014, 40(4):1-6.
[12] 庄逢辰.液体火箭发动机喷雾燃烧的理论、模型及应用[M].长沙:国防科技大学出版社, 1995.
[13] 王尧, 李国岫, 虞育松, 等.喷射压力对同轴旋转射流喷雾锥角影响的实验研究[J].推进技术,2017,38(4):188-193.
[14] 潘华辰, 周泽磊, 刘雷.关键结构参数对离心式雾化喷嘴雾化效果的影响研究[J].机械工程学报, 2017, 53(2):199-206.
[15] 李进贤, 何浩波, 岳春国, 等.旋流式气/液同轴喷嘴常压雾化燃烧实验研究[J].宇航学报, 2008, 29(5):1563-1569.
[16] TRINKS H, HOFFMAN R.Experimental investigation of bipropellant exhaust plume flowfield, heating, and contamination, and comparison with the CONTAM computer model predictions[C]//18th Thermophysics Conference. Montreal, Canada,Reston, Virigina:AIAA, 1983.
[17] 杨成虎, 刘犇.喷雾场测试技术研究进展[J].火箭推进, 2010, 36(4):16-23.YANG C H, LIU B.Development of diagnostic techniques for spray measurements[J].Journal of Rocket Propulsion, 2010, 36(4):16-23.
[18] 张蒙正, 傅永贵, 张泽平, 等.两股互击式喷嘴雾化研究及应用[J].推进技术, 1999, 20(2):19-26.
[19] 周红玲.液体远地点发动机优化设计及系统动态性能研究[D].长沙:国防科学技术大学, 2017.
[20] 张绪虎, 汪翔, 贾中华, 等.小推力姿控轨控火箭发动机材料技术研究现状[J].导弹与航天运载技术, 2005(6):32-37.
备注/Memo
收稿日期:2019-07-03; 修回日期:2019-10-11基金项目:国家装备预先研究项目(30508040301)作者简介:赵婷(1984—),女,高级工程师,研究领域为液体火箭发动机设计
更新日期/Last Update:
2020-02-25