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首页> 外国专利> METHOD attitude control orbiting spacecraft with inertia of the executive bodies in probing the Earth's atmosphere

METHOD attitude control orbiting spacecraft with inertia of the executive bodies in probing the Earth's atmosphere

机译：方法用姿态执行器惯性控制航天器，探测地球大气

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摘要

method u0443u043fu0440u0430u0432u043bu0435u043du0438u00a0 orientation of orbital spacecraft with inertia executive bodies by sensing the earth's atmosphere, including u043du0430u0432u0438u0433u0430u0446 ionic u0438u0437u043cu0435u0440u0435u043du0438u00a0, a spacecraft to u0441u043eu0432u043cu0435u0449u0435u043du0438u00a0 axis of minimum moment of inertia of the spacecraft in the orbit plane and the subsequent u0441u0442u0430u0431u0438u043bu0438 held spacecraft in inertial coordinate system, u043eu0442u043bu0438u0447u0430u044eu0449u0438u0439u0441u00a0 orderfurther u0438u0437u043cu0435u0440u00a0u044eu0442 orbital altitude of the spacecraft and u043eu043fu0440u0435u0434u0435u043bu00a0u044eu0442 on her u0437u043du0430u0447u0435u043du0438u00a0 of u043fu043eu043bu0443u0440u0430u0441u0442u0432u043eu0440u0430 visible from the spacecraft earth (disk q) and the upper border of the u0432u043eu0437u0432u044bu0448u0435u043du0438u00a0 u0441u043bu043eu00a0 atmosphere on the visible horizon with spacecraft earth (), u0438u0437u043cu0435u0440u00a0u044eu0442 u0443u0433u043bu043eu0432u0443u044e orbital speed d. u0432u0438u0436u0435u043du0438u00a0 spacecraft ()u0432u044bu043fu043eu043bu043du00a0u044eu0442 turnaround spacecraft around the axis of minimum moment of inertia to u0441u043eu0432u043cu0435u0449u0435u043du0438u00a0 axis u0432u0438u0437u0438u0440u043eu0432u0430u043du0438u00a0 device u0437u043eu043du0434u0438u0440u043eu0432u0430u043du0438u00a0 with plane orbit u043au043eu0441u043cu0438u0447 u0435u0441u043au043eu0433u043e apparatus, the axis of minimum moment of inertia with the axis of the device which u0441u043eu0441u0442u0430u0432u043bu00a0u0435u0442 u0432u0438u0437u0438u0440u043eu0432u0430u043du0438u00a0 u0437u043eu043du0434u0438u0440u043eu0432u0430u043du0438u00a0 angle greater than or equal to the value of q, and follow a crucial stabilizing the spacecraft in inertial coordinate system.fixed time (tO) u0441u043eu0432u043cu0435u0449u0435u043du0438u00a0 u043du0430u043fu0440u0430u0432u043bu0435u043du0438u00a0 axis u0432u0438u0437u0438u0440u043eu0432u0430u043du0438u00a0 device u0437u043eu043du0434u0438u0440u043eu0432u0430u043du0438u00a0 direction, opposite to the speed vector of the spacecraft provided that u0444u0438u043au0441u0438u0440u043eu0432u0430u043du043du0430u00a0 axis of minimum moment of inertia of the spacecraft at the same time u0441u043eu0441u0442u0430u0432u043bu00a0u0435u0442 u043fu0440u00a0u043cu044bu0435 or sharp edges with a radius vector space the apparatus and the vector.the opposite vector velocity of the spacecraft, the angle between the axis of the device continuously u0438u0437u043cu0435u0440u00a0u044eu0442 u0432u0438u0437u0438u0440u043eu0432u0430u043du0438u00a0 u0437u043eu043du0434u0438u0440u043eu0432u0430u043du0438u00a0 and focus on visible with u043au043eu0441u043cu0438u0447 the apparatus u0435u0441u043au043eu0433u043e horizon land () and at the time of his u043eu0431u043du0443u043bu0435u043du0438u00a0 spread the spacecraft around the axis of minimum moment of inertia to u0434u043eu0441u0442u0438u0436u0435u043du0438u00a0 u0437u043du0430u0447u0435u043du0438u00a0 angle equal value angleby u0432u044bu043fu043eu043bu043du0435u043du0438u00a0 turnaround spacecraft in the orientation obtained from the angular u043fu043eu043bu043eu0436u0435u043du0438u00a0 spacecraft at the present time tO cook autumn spacecraft around a fixed axis of minimum moment of inertia at the corner, u043eu043fu0440u0435u0434u0435u043bu00a0u0435u043cu044bu0439 formula;,;where tk is the time u043eu043au043eu043du0447u0430u043du0438u00a0 the u-turn;- the angle between the axis of the device u0432u0438u0437u0438u0440u043eu0432u0430u043du0438u00a0 u0437u043eu043du0434u0438u0440u043eu0432u0430u043du0438u00a0 and axis of minimum moment of inertia of the spacecraft.;then continue to stabilize the spacecraft in inertial coordinate system, u043eu043fu0440u0435u0434u0435u043bu00a0u044eu0442 moment u043eu0431u043du0443u043bu0435u043du0438u00a0 u0438u0437u043cu0435u0440u00a0u0435u043cu043eu0433u043e angle and at the time of his u043eu0431u043du0443u043bu0435u043du0438 u00a0 u043fu043eu0432u0442u043eu0440u00a0u044eu0442 the u0434u0435u0439u0441u0442u0432u0438u00a0, and after the point of time u043fu0440u043eu0445u043eu0436u0434u0435u043du0438u00a0 u043fu043eu043bu043eu0436u0435u043du0438u00a0 in orbitwhich u043fu0440u043eu0435u043au0446u0438u00a0 axis u0432u0438u0437u0438u0440u043eu0432u0430u043du0438u00a0 device u0437u043eu043du0434u0438u0440u043eu0432u0430u043du0438u00a0 on plane of orbit spacecraft is directed against the radius vector of the spacecraft, nast u043fu0430u044eu0449u0435u0433u043e after a time interval t from the time length of the recorded tO, where the t u043eu043fu0440u0435u0434u0435u043bu00a0u0435u0442u0441u00a0 in formula;,;and in the formula;,;compare the value with the value of the u0438u0437u043cu0435u0440u00a0u0435u043cu043eu0433u043e corner angle and moment equality spread data of spacecraft around the axis of minimum moment of yin u0440u0446u0438u0438 to u0434u043eu0441u0442u0438u0436u0435u043du0438u00a0 zero u0437u043du0430u0447u0435u043du0438u00a0 angle by u0432u044bu043fu043eu043bu043du0435u043du0438u00a0 turnaround spacecraft in orientation.derived from the u043fu043eu043bu043eu0436u0435u043du0438u00a0 spacecraft at the present time tO rotates around a fixed axis of a spacecraft the moment of inertia at the corner, u043eu043fu0440u0435u0434u0435u043bu00a0u0435u043cu044bu0439 in formula;,;and in the formula;,;where tk is the time u043eu043au043eu043du0447u0430u043du0438u00a0 the u-turn, then continue to stabilize the spacecraft in inertial coordinate system, compare the u0437u043du0430u0447u0435u043du0438 (e u0438u0437u043cu0435u0440u00a0u0435u043cu043eu0433u043e angle with the value of the angle and time data of the u0434u0435u0439u0441u0442u0432u0438u00a0 u043fu043eu0432u0442u043eu0440u00a0u044eu0442 equality.
机译：通过感应地球大气层，包括惯性执行体的方法，将 u0443 u043f u0440 u0430 u0432 u043b u0435 u043d u0438 u00a0定向，包括 u043d u0430 u0432 u0433 u0433 ionic u0438 u0437 u043c u0435 u0440 u0435 u043d u0438 u00a0，这是航天器到 u0441 u043e u0432 u043c u0435 u0449 u0435 u043d u043d u0438 u00a0的最小惯性矩轴航天器在轨道平面上以及随后的 u0441 u0442 u0430 u0431 u0438 u043b u0438将航天器保持在惯性坐标系中， u043e u0442 u043b u0438 u0447 u0430 u044e u0449 u0438 u0439 u0441 u00a0进一步 u0438 u0437 u043c u0435 u0440 u00a0 u044e u0442航天器的轨道高度和 u043e u043f u0440 u0435 u0434 u0435 u0435 u043b u00a0 u044e u0442 on从航天器地球（磁盘q）可见的u0437 u043d u0430 u0447 u0435 u043d u0438 u00a0 u043f u043e u043b u0443 u0440 u0430 u0441 u0442 u0432 u043e u0440 u0440 u0430和 u0432 u043e u0437 u043的上边界2 u044b u0448 u0435 u043d u0438 u00a0 u0441 u043b u043e u00a0可见地平线上的大气层与航天器地球（）， u0438 u0437 u043c u0435 u0440 u00a0 u044e u0442 u0443 u0433 u043b u043e u0432 u0443 u044e轨道速度d。 u0432 u0438 u0436 u0435 u043d u0438 u00a0（） u0432 u044b u043f u043e u043b u043d u043d u00a0 u044e u0442围绕最小惯性矩轴将飞行器转至 u0441 u043e u0432 u043c u0435 u0449 u0435 u043d u0438 u00a0轴 u0432 u0438 u0437 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0设备 u0437 u043e u043d u0434 u0440 u043e u0432 u0430 u043d u0438 u00a0具有平面轨道 u043a u043e u0441 u043c u0438 u0447 u0435 u0441 u043a u043e u043e u0433 u043e装置，最小惯性矩的轴 u0441 u043e u0441 u0442 u0430 u0432 u043b u00a0 u0435 u0442 u0432 u0438 u0437 u0438 u0440 u043e u043e u0432 u0432 u0430 u043d u0438 u004380的设备的轴u0437 u043e u043d u0434 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0角度大于或等于q的值，并遵循在惯性坐标系中稳定航天器的关键步骤。固定时间（tO ） u0441 u043e u0432 u043c u0435 u0449 u0435 u043d u0438 u00a0 u043d u0430 u043f u0440 u043 0 u0432 u043b u0435 u043d u0438 u00a0轴 u0432 u0438 u0437 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0设备 u0437 u043e u043d u0434 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0方向，与航天器的速度矢量相反，但前提是 u0444 u0438 u043a u0441 u0438 u0440 u043e u0432 u0430 u043d u043d u043d u0430 u00a0轴 u0441 u043e u0441 u0442 u0430 u0432 u043b u00a0 u0435 u0442 u043f u0440 u00a0 u043c u044c u044b u0440 u00a0 u043c u044b u0435或半径较大的锐边向量将设备和向量间隔开。与航天器相反的向量速度，设备轴之间的角度连续 u0438 u0437 u043c u0435 u0440 u00a0 u044e u0442 u0432 u0438 u0437 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0 u0437 u043e u043d u043d u0434 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0并通过 u043a u043e u0431使其可见 u0438 u0447设备 u0435 u0441 u043a u043e u0433 u043e地平线土地（）并且在他的 u043e u0431 u043d u0443 u043b u0435 u043d u0438 u00a0时，将航天器围绕最小惯性矩轴展开到 u0434 u043e u0441 u0442 u0438 u0436 u0435 u043d u0438 u00a0 u0437 u043d u0430 u0447 u0435 u043d u0438 u00a0角度相等值angle by u0432 u044b u043f u043e u043b u043d u0435 u043d u043d u0438 u00a0从当前的角度 u043f u043e u043b u043e u0436 u0435 u043d u0438 u00a0获得的方向，以秋天的航天器围绕转角最小惯性矩的固定轴 u043e u043f u0440 u0435 u0434 u0435 u043b u00a0 u0435 u043c u044b u0439公式;;其中tk是时间 u043e u043a u043e u043d u0447 u0430 u043d u0438 u00a0 ;-设备轴之间的角度 u0432 u0438 u0437 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0 u0437 u043e u043d u0434 u0438 u0440 u043e u0432 u0432 u043d u0438 u00a0和航天器的最小惯性矩轴。继续在惯性坐标系中稳定航天器， u043e u043f u0440 u0435 u0434 u0435 u043b u00a0 u044e u0442矩 u043e u0431 u043d u0443 u043b u0435 u043d u043d u0438 u00a0 u0438 u0437 u043c u0435 u0440 u00a0 u0435 u043c u043e u0433 u043e角度以及他的 u043e u0431 u043d u0443 u043b u0435 u043d u043d u0438 u00a0 u043f u043e u0432 u0442 u043e u0440 u00a0 u044e u0442 u0434 u0435 u0439 u0441 u0442 u0432 u0438 u00a0，并且在时间点 u043f u0440 u043e u0445 u043e u0436 u0434 u0435 u043d u043d u0438 u00a0 u043f u043e u043b u043e u0436 u0435 u0435 u043d u043d u00438 u00a0 u043f u0440 u043e u0435 u043a u0446 u0438 u00a0轴 u0432 u0438 u0437 u0438 u0440 u043e u0432 u0430 u043d u0438 u00a0设备 u0437 u043e u043d u0434距航天器平面u0440 u043e u0432 u0430 u043d u0438 u00a0指向航天器的半径矢量，nast u043f u0430 u044e u0449 u0435 u0433 u043e记录的tO的时间长度，其中t u043e u043f u0440 u0435 u0434 u0435 u043b u00a0 u0435 u0442 u0441 u00a0在公式中;将公式中的值与 u0438 u0437 u043c u0435 u0440 u00a0 u0435 u043c u043e u0433 u043e的角和矩等距值绕航天器的最小弯矩 u0440 u0446 u0438 u0438至 u0434 u043e u0441 u0442 u0438 u0436 u0435 u043d u0438 u00a0零 u0437 u043d u由 u0432 u044b u043b u043f u043e u043b u043d u0435 u043d u0438 u00a0转向的飞行角度为0430 u0447 u0435 u043d u0438 u00a0角度。来自 u043f u043e u043b u043e u0436 u0435 u043d u0438 u00a0当前时刻tO会在拐角处的惯性矩 u043e u043f u0440 u0435 u0434 u0435 u043b u00a0 u0435 公式中，其中tk是U形转弯的时间 u043e u043a u043e u043d u0447 u0430 u043d u0438 u00a0然后继续稳定航天器在惯性坐标系中，将 u0437 u043d u0430 u0447 u0435 u043d u0438（e u0438 u0437 u043c u0435 u0440 u00a0 u0435 u043c u043e u0433 u043e角度与该值进行比较 u0434 u0435 u0439 u0441 u0442 u0432 u0438 u00a0 u043f u043e u0432 u0442 u043e u0440 u00a0 u044e u0442的角度和时间数据。

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