The invention relates to thermal control facilities of attitude and orbit control engines of aerospace crafts, in particular to a thrust chamber integrated thermal control device for a micro space liquefied gas thruster. An armor shell of the device is provided with an armor body, the armor body is a hollow cylinder and comprises a working section, a leading-out section and a welding table which share the same axis, center holes of the working section, the leading-out section and the welding table are communicated to form a middle cavity, and a heat storage and exchange device is arranged in the middle cavity; mounting grooves or mounting holes are uniformly processed in the armor wall of the working section along the circumferential direction, the leading-out section is provided with an annular leading-out groove, and leading-out holes are uniformly processed in the outer wall of the leading-out section along the circumferential direction; a heating core is fixed in the mounting groove or the mounting hole of the armor body working section; and the heating core is connected with the outer lead in a transition lead-out mode of a heating core transition lead-out assembly. The device is applied to heat control of a thrust chamber of an aerospace craft propelling system. The heat storage and exchange performance, the electric heating performance, the insulating performance and the assembling performance of the device all exceed the using requirements.
本发明涉及航天飞行器姿、轨控发动机的热控设施,具体是一种微型空间液化气推力器用推力室一体化热控装置。该装置的铠壳设有铠体,铠体为中空的圆柱体,包括同一轴线的工作段、引出段和焊台,工作段、引出段和焊台的中心孔相连通形成中部空腔,所述中部空腔内设置蓄换热器;工作段的铠壁中沿圆周方向均匀加工安装槽或安装孔,引出段设有环形引出槽,其外壁上沿圆周方向均匀加工引出孔;发热芯固定于铠体工作段的安装槽或安装孔中,发热芯与外引线的连接采用发热芯过渡引出组件过渡引出方式。本发明应用于航天飞行器推进系统推力室的热控,其蓄换热性能、电热性能、绝缘性能和装配性能均超过使用要求。
Thrust chamber integrated thermal control device for micro space liquefied gas thruster
一种微型空间液化气推力器用推力室一体化热控装置
2021-12-03
Patent
Electronic Resource
Chinese
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