WO2021102669A1 - 超低轨道卫星轨道自主维持方法 - Google Patents

超低轨道卫星轨道自主维持方法 Download PDF

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WO2021102669A1
WO2021102669A1 PCT/CN2019/120878 CN2019120878W WO2021102669A1 WO 2021102669 A1 WO2021102669 A1 WO 2021102669A1 CN 2019120878 W CN2019120878 W CN 2019120878W WO 2021102669 A1 WO2021102669 A1 WO 2021102669A1
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ultra
orbit
low
satellite
low orbit
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PCT/CN2019/120878
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English (en)
French (fr)
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刘国华
程蛟
姚小松
戴正升
祁海铭
田龙飞
范城城
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中国科学院微小卫星创新研究院
上海微小卫星工程中心
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Priority to CN201980020437.2A priority Critical patent/CN111989265B/zh
Priority to PCT/CN2019/120878 priority patent/WO2021102669A1/zh
Publication of WO2021102669A1 publication Critical patent/WO2021102669A1/zh

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B64G1/245—Attitude control algorithms for spacecraft attitude control
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation

Definitions

  • the invention relates to the field of aerospace technology, in particular to a method for autonomously maintaining the orbit of an ultra-low orbit satellite.
  • Ultra-low orbit refers to an orbit that is flying outside the atmosphere but below the orbital height of an ordinary spacecraft. It usually refers to a flight orbit with an orbit height above 120km and below 300km. Ultra-low orbit has great military and scientific significance. However, due to the low orbital altitude, atmospheric drag has a significant impact on the orbit when flying in ultra-low orbit. If the orbit is not maintained, the satellite's orbit will rapidly decay.
  • ultra-low orbit satellites have long-term operation at 200-300km.
  • the atmosphere is dense, the aircraft can maintain the perturbation of atmospheric drag by maintaining high-frequency orbit.
  • Long time operation in orbit taking into account the advantages of nearby spacecraft and orbiting satellites.
  • the commonly used orbit maintenance method is mainly based on the orbit measurement of a certain frequency.
  • the orbit correction is carried out through the closed loop of the sky and the earth or the closed loop on the satellite.
  • the aerodynamic moment is greatly increased, the orbit attenuation increases, and the number of orbit maintenance increases accordingly;
  • the ultra-low orbit satellite orbit maintenance method adopts high-thrust orbit control, and the strong interference introduced during the high-thrust orbit control brings about the availability of satellites.
  • the orbit maintenance accuracy is related to the maintenance frequency. Combined with the current orbit measurement accuracy and the thruster's working characteristics, the orbit maintenance accuracy is on the order of km.
  • the orbit maintenance accuracy of ultra-low orbit satellites is relatively low at this stage.
  • the purpose of the present invention is to provide a method for autonomously maintaining the orbit of an ultra-low orbit satellite to solve the problem of low orbit maintenance accuracy of the existing ultra-low orbit satellite.
  • the present invention provides a method for autonomously maintaining the orbit of an ultra-low orbit satellite.
  • the method for autonomously maintaining the orbit of an ultra-low orbit satellite includes:
  • Step 1 Set the working orbit range of the ultra-low orbit satellite and estimate the magnitude of the atmospheric resistance
  • Step 2 According to the magnitude of the atmospheric resistance, analyze the magnitude of the noise of the inertial acceleration measurement system to obtain the noise analysis result of the inertial acceleration measurement system;
  • Step 3 According to the noise analysis result of the inertial acceleration measurement system, set the parameters of the small thrust execution system, and perform on-orbit calibration of the inertial acceleration measurement system and the small thrust execution system to obtain the calibrated inertial acceleration output result ;as well as
  • Step 4 According to the calibrated inertial acceleration output result, the orbit control small thrust output algorithm of the small thrust execution system is set.
  • the inertial acceleration measurement system includes an inertial sensor, and the inertial sensor is used to measure the acceleration of the ultra-low orbit satellite; and/or
  • the small thrust execution system includes a thruster, and the thruster is used to provide power for the ultra-low orbit satellite.
  • the first step includes:
  • the value of the atmospheric resistance is determined according to the atmospheric density and the direction of the incoming flow.
  • the step one further includes:
  • F is the value of the atmospheric drag
  • C d is the drag coefficient
  • ⁇ is the atmospheric density within the working orbit range
  • S is the windward area
  • v is the speed of the atmosphere relative to the satellite
  • the magnitude of the atmospheric resistance is obtained.
  • the step one further includes:
  • the second step includes:
  • the noise of the inertial acceleration measurement system includes the disturbance of the spacecraft by the acceleration of the atmospheric drag, and a number of other disturbances;
  • the proportion of the atmospheric drag acceleration disturbance of the spacecraft in the noise of the inertial acceleration measurement system is evaluated, so that the magnitude of each of the other disturbances is greater than that of the spacecraft
  • the magnitude of disturbance by atmospheric drag acceleration is one level lower.
  • the second step further includes:
  • F is the value of the atmospheric resistance
  • m is the mass of the ultra-low orbit satellite.
  • the other disturbances include at least one of the following: residual acceleration noise of the inertial sensor, direct acceleration disturbance of the test mass, coupling stiffness error between the spacecraft and the test mass, and the sum of high-frequency noise and quantization noise; and
  • the second step further includes: acquiring the accuracy of the inertial acceleration measurement system, and acquiring at least one of the following items according to the accuracy of the inertial acceleration measurement system: residual acceleration noise of the inertial sensor, direct acceleration of the test mass Disturbance, the coupling stiffness error of the spacecraft and the test mass, and the sum of the high-frequency noise and the quantization noise.
  • the step three includes:
  • the parameters of the small thrust execution system include at least one of the following items: the control frequency of the small thrust execution system, the continuous working time of the thruster, and the thruster design margin of the small thrust execution system , Determining the thruster design margin of the small thrust execution system according to the control frequency and the continuous working time of the thruster;
  • f is the thrust
  • K is the design margin of the thruster
  • ⁇ is the noise of the inertial acceleration measurement system.
  • the step three further includes:
  • Performing on-orbit calibration of the inertial acceleration measurement system to obtain the calibrated inertial acceleration output result includes: calibrating the inertial sensor, setting the attitude of the ultra-low orbit satellite, and the inertial sensor measuring the ultra-low orbit multiple times. Obtain multiple first acceleration measurement values from the acceleration of the low-orbit satellite, and calculate an average value of the multiple first acceleration measurement values, and the average value is used as the calibrated inertial acceleration output result;
  • the on-orbit calibration of the small thrust execution system includes: after completing the on-orbit calibration of the inertial acceleration measurement system, calibrating the thruster, the thruster is ignited once, and the inertial sensor measures the ultra-low Obtain the second acceleration measurement value from the acceleration of the orbiting satellite, and calculate the state quantity of each control period of the thruster output result according to the mass of the ultra-low orbit satellite and the plurality of second acceleration measurement values.
  • the step three includes:
  • the number of times that the inertial sensor measures the acceleration of the ultra-low orbit satellite is greater than 100 times;
  • the small thrust is The execution system performs on-orbit calibration.
  • the step four includes:
  • X k is the state quantity of the kth control cycle of the output result of the thruster
  • a ob is the thruster installation matrix
  • B ob is the no-drag torque output matrix
  • F k is the drag-free control thrust applied in the k-th control cycle
  • Acc K+1 is the preprocessed relative acceleration measured by the inertial sensor in the k+1 control cycle
  • m is the satellite quality.
  • step five which includes:
  • the non-conservative force in the ultra-low orbit environment dominated by aerodynamic moment is calculated.
  • the magnitude of the atmospheric resistance is estimated by setting the working orbit range of the satellite, the magnitude of the noise of the inertial acceleration measurement system is analyzed, and the noise analysis result of the inertial acceleration measurement system is used. Set the parameters of the small thrust execution system, and perform on-orbit calibration of the inertial acceleration measurement system and the small thrust execution system.
  • the satellite orbit autonomous maintenance method based on continuous small thrust meets the satellite's long-term high-precision maintenance requirements for ultra-low orbits, and satisfies the long-term autonomous maintenance needs of satellites operating in ultra-low orbits between 200 and 300 km.
  • the control frequency of the small thrust execution system and the continuous working time of the thruster of the small thrust execution system are determined according to the satellite orbit control strategy, and the thruster design margin of the small thrust execution system is determined according to the control frequency and the continuous working time of the thruster.
  • the thrust of the small thrust execution system is set, and the inertial acceleration measurement system is calibrated on orbit, which realizes the ultra-low orbit environment dominated by the aerodynamic moment based on the relative acceleration provided by the inertial sensor.
  • an orbit maintenance algorithm based on continuous small thrust is designed.
  • the orbit maintenance accuracy is on the order of 200m, which meets the long-term autonomous maintenance needs of ultra-low orbit satellites and improves the orbit maintenance accuracy.
  • the method for autonomous maintenance of the ultra-low orbit satellite orbit of the present invention has strong theoretical application and engineering practical value, and has obvious effects. Since the software is mainly improved, the hardware does not need to be replaced, which is convenient for engineering realization.
  • Fig. 1 is a schematic flowchart of a method for autonomously maintaining the orbit of an ultra-low orbit satellite according to an embodiment of the present invention.
  • the core idea of the present invention is to provide a method for autonomously maintaining the orbit of an ultra-low orbit satellite to solve the problem of low orbit maintenance accuracy of the existing ultra-low orbit satellite.
  • the present invention provides a method for autonomously maintaining the orbit of an ultra-low orbit satellite.
  • the method for autonomously maintaining the orbit of an ultra-low orbit satellite includes: step one: setting the operating orbit range of the satellite and estimating the magnitude of the atmospheric resistance; Two: According to the magnitude of the atmospheric resistance, analyze the magnitude of the noise of the inertial acceleration measurement system to obtain the noise analysis result of the inertial acceleration measurement system; Step 3: Set the small thrust execution system according to the noise analysis result of the inertial acceleration measurement system And perform on-orbit calibration of the inertial acceleration measurement system and the small thrust execution system to obtain the calibrated inertial acceleration output result; and step 4: according to the calibrated inertial acceleration output result, set the The orbit control small thrust output algorithm of the small thrust execution system.
  • the method for autonomously maintaining the orbit of an ultra-low orbit satellite includes: Step 2: According to the magnitude of the atmospheric resistance, analyze the magnitude of the noise of the inertial acceleration measurement system to obtain the noise analysis result of the inertial acceleration measurement system; Step 3: Set according to the noise analysis result of the inertial acceleration measurement system The parameters of the small thrust execution system, and perform on-orbit calibration of the inertial acceleration measurement system and the small thrust execution system to obtain the calibrated inertial acceleration output result; and step 4: output the result according to the calibrated inertial acceleration , Set the orbit control small thrust output algorithm of the small thrust execution system.
  • the inertial acceleration measurement system includes an inertial sensor, and the inertial sensor is used to measure the acceleration of the ultra-low orbit satellite; and/or the small thrust
  • the execution system includes a thruster, and the thruster is used to provide the power of the ultra-low orbit satellite.
  • the first step includes: obtaining the configuration of the ultra-low orbit satellite; setting the working orbit range of the ultra-low orbit satellite; After the configuration of and the range of the working orbit are determined, the value of the atmospheric resistance is determined according to the atmospheric density and the direction of the incoming flow.
  • the first step further includes: obtaining the windward area and the relative atmospheric pressure according to the configuration of the ultra-low orbit satellite and the working orbit range of the ultra-low orbit satellite.
  • the speed of the ultra-low orbit satellite obtain the atmospheric density within the operating orbit range according to the operating orbit range; obtain the operating direction of the ultra-low orbit satellite on the operating orbit, and calculate according to the operating direction
  • the direction of the incoming flow the resistance coefficient is calculated according to the atmospheric density and the direction of the incoming flow
  • the value of the atmospheric resistance is estimated according to the above-mentioned parameters:
  • the step one further includes: calculating the changing trend of the incoming flow direction within one orbit period of the ultra-low orbit satellite's ultra-low orbit operation, according to the The change trend of the incoming flow direction in one orbit period of the ultra-low orbit operation of the ultra-low orbit satellite is estimated to estimate the average atmospheric drag during one orbit period of the ultra-low orbit satellite's ultra-low orbit operation.
  • the second step includes: the noise of the inertial acceleration measurement system includes the acceleration disturbance of the spacecraft by the atmospheric drag, and a number of other disturbances; the measurement of the inertial acceleration The noise of the system is decomposed, and it is obtained that the spacecraft is disturbed by the atmospheric drag acceleration; according to the disturbance of the spacecraft by the atmospheric drag acceleration, it is estimated that the spacecraft is affected by the atmospheric drag acceleration disturbance in the noise of the inertial acceleration measurement system.
  • the specific gravity is such that the magnitude of each of the other disturbances is one level lower than the magnitude of the spacecraft disturbed by the atmospheric drag acceleration.
  • the second step further includes: obtaining the mass of the ultra-low orbit satellite, and calculating aerospace based on the value of the atmospheric drag and the mass of the ultra-low orbit satellite The device is disturbed by the acceleration of aerodynamic drag:
  • the other disturbances include at least one of the following: residual acceleration noise of the inertial sensor, direct acceleration disturbance of the test mass, and coupling stiffness between the spacecraft and the test mass Error, and the sum of high-frequency noise and quantization noise; and the second step further includes obtaining the accuracy of the inertial acceleration measurement system, and obtaining at least one of the following items according to the accuracy of the inertial acceleration measurement system: the inertial sensor The residual acceleration noise of the test mass, the direct acceleration disturbance of the test mass, the coupling stiffness error of the spacecraft and the test mass, and the sum of the high-frequency noise and the quantization noise.
  • the step three includes: setting the parameters of the low-thrust execution system, wherein the parameters of the low-thrust execution system include at least one of the following items:
  • the control frequency of the thrust execution system, the continuous working time of the thruster, and the thruster design margin of the small thrust execution system, and the thruster design margin of the small thrust execution system is determined according to the control frequency and the continuous working time of the thruster Degree; according to the noise analysis result of the inertial acceleration measurement system, set the thrust of the small thrust execution system:
  • the step three further includes: calibrating the inertial acceleration measurement system on orbit, and obtaining the calibrated inertial acceleration output result includes: calibrating the inertial sensor , Setting the attitude of the ultra-low orbit satellite, the inertial sensor measures the acceleration of the ultra-low orbit satellite multiple times to obtain a plurality of first acceleration measurement values, and calculate the average value of the plurality of first acceleration measurement values, The average value is used as the calibrated inertial acceleration output result;
  • the on-orbit calibration of the small thrust execution system includes: calibrating the thruster after completing the on-orbit calibration of the inertial acceleration measurement system, The thruster ignites once, the inertial sensor measures the acceleration of the ultra-low orbit satellite to obtain a second acceleration measurement value, and calculates the second acceleration measurement value based on the mass of the ultra-low orbit satellite and the multiple second acceleration measurement values
  • the step three includes: performing on-orbit calibration of the inertial acceleration measurement system, and the number of times the inertial sensor measures the acceleration of the ultra-low orbit satellite is greater than 100 Times; before the inertial acceleration measurement system is calibrated in orbit, the attitude of the ultra-low orbit satellite is adjusted to within the required accuracy of the index, so as to reduce the attitude change of the ultra-low orbit satellite within the calibration time, so as to reduce the ultra-low orbit satellite
  • the projection component of the external interference in the direction of the sensitive axis caused by the attitude change of the low-orbit satellite when the inertial acceleration measurement system is on-orbit calibration, the ultra-low orbit satellite is set in the attitude free drift mode to reduce the ultra-low orbit Coupling interference caused by satellite attitude control to inertial sensor measurement calibration; according to the change trend of the incoming flow direction within one orbit period of the ultra-low orbit satellite operation and the ultra-low orbit satellite operation
  • the average atmospheric drag in an orbital period performs on-orbit calibration
  • the step four includes: setting the orbit control small thrust output algorithm of the small thrust execution system according to the calibrated inertial acceleration output result, and calculating the current The output result of the thruster of the small thrust execution system at all times, and closed-loop control:
  • X k is the state quantity of the k-th control cycle of the thruster output result
  • a ob is the thruster installation matrix
  • Lob is the installation matrix of the inertial acceleration measurement system
  • B ob is the no-drag torque output matrix
  • F k is the drag-free control thrust applied in the k-th control period
  • Acc K+1 is the preprocessed relative acceleration measured by the inertial sensor in the k+1-th control period
  • m is the mass of the satellite.
  • the method for autonomously maintaining the orbit of the ultra-low orbit satellite further includes step five, and the step five includes: according to the preprocessed relative value measured by the inertial sensor. Acceleration, calculate the non-conservative force in the ultra-low orbit environment dominated by aerodynamic moment.
  • the magnitude of the atmospheric resistance is estimated by setting the working orbit range of the satellite, the magnitude of the noise of the inertial acceleration measurement system is analyzed, and the noise analysis result of the inertial acceleration measurement system is used. Set the parameters of the small thrust execution system, and perform on-orbit calibration of the inertial acceleration measurement system and the small thrust execution system.
  • the satellite orbit autonomous maintenance method based on continuous small thrust meets the satellite's long-term high-precision maintenance requirements for ultra-low orbits, and satisfies the long-term autonomous maintenance needs of satellites operating in ultra-low orbits between 200 and 300 km.
  • the control frequency of the small thrust execution system and the continuous working time of the thruster of the small thrust execution system are determined according to the satellite orbit control strategy, and the thruster design margin of the small thrust execution system is determined according to the control frequency and the continuous working time of the thruster.
  • the thrust of the small thrust execution system is set, and the inertial acceleration measurement system is calibrated on orbit, which realizes the ultra-low orbit environment dominated by the aerodynamic moment based on the relative acceleration provided by the inertial sensor.
  • an orbit maintenance algorithm based on continuous small thrust is designed.
  • the orbit maintenance accuracy is on the order of 200m, which meets the long-term autonomous maintenance needs of ultra-low orbit satellites and improves the orbit maintenance accuracy.
  • the method for autonomous maintenance of the ultra-low orbit satellite orbit of the present invention has strong theoretical application and engineering practical value, and has obvious effects. Since the software is mainly improved, the hardware does not need to be replaced, which is convenient for engineering realization.
  • the above embodiments describe in detail the different configurations of the method for autonomous maintenance of the ultra-low orbit satellite orbit.
  • the present invention includes but is not limited to the configurations listed in the above implementations, and any configuration provided in the above embodiments.
  • the content that is transformed on the basis of the model falls within the protection scope of the present invention.
  • Those skilled in the art can draw inferences based on the content of the above-mentioned embodiments.

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Abstract

一种超低轨道卫星轨道自主维持方法，该方法包括：步骤一、设置卫星的工作轨道范围，估算大气阻力的量级；步骤二：根据大气阻力的量级，分析惯性加速度测量系统的噪声的量级，得到惯性加速度测量系统噪声分析结果；步骤三：根据惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对惯性加速度测量系统和小推力执行系统进行在轨标定，得到标定后的惯性加速度输出结果；以及步骤四：根据标定后的惯性加速度输出结果，设置小推力执行系统的轨控小推力输出算法。

Description

超低轨道卫星轨道自主维持方法 技术领域
本发明涉及航空航天技术领域，特别涉及一种超低轨道卫星轨道自主维持方法。
背景技术
超低轨道是指飞行在大气层外，但低于普通航天器轨道高度的轨道，通常指轨道高度在120km以上，300km以下的飞行轨道。超低轨道具有较大的军事和科学意义，但由于轨道高度低，在超低轨道飞行时，大气阻力对轨道影响显著，若不进行轨道维持，卫星轨道会迅速衰减。
随着现代卫星技术的不断发展，超低轨卫星长期运行在200～300km，在该区域，虽然大气密度较大，但飞行器通过持续高频度轨道维持，可以将大气阻力的摄动衰减作用抵消，在轨较长时间运行，兼顾临近空间飞行器和轨道卫星的优势。
目前常用的轨道维持方法主要是基于一定频度的轨道测量，根据轨道衰减量，通过天地闭环或星上闭环，采用大推力进行轨道修正，使得超低轨卫星所处的空间环境与传统轨道卫星相比，气动力矩大大增加，轨道衰减增加，轨道维持的次数也相应增加；另外，超低轨卫星轨道维持方法采用大推力轨道控制，在大推力轨控期间引入的强干扰对于卫星可用性带来了一定的影响；进一步的，轨道维持精度和维持频率相关，结合目前的轨道测量精度和推力器的工作特性，轨道维持精度在km量级。
综上所述，现阶段超低轨卫星的轨道维持精度较低。
发明内容
本发明的目的在于提供一种超低轨道卫星轨道自主维持方法，以解决现有的超低轨卫星的轨道维持精度较低的问题。
为解决上述技术问题，本发明提供一种超低轨道卫星轨道自主维持方法，所述超低轨道卫星轨道自主维持方法包括：
步骤一、设置超低轨道卫星的工作轨道范围，估算大气阻力的量级；
步骤二：根据所述大气阻力的量级，分析惯性加速度测量系统的噪声的量级，得到惯性加速度测量系统噪声分析结果；
步骤三：根据所述惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对所述惯性加速度测量系统和所述小推力执行系统进行在轨标定，得到标定后的惯性加速度输出结果；以及
步骤四：根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述惯性加速度测量系统包括惯性传感器，所述惯性传感器用于测量所述超低轨道卫星的加速度；和/或
所述小推力执行系统包括推力器，所述推力器用于提供所述超低轨道卫星的动力。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤一包括：
获取所述超低轨道卫星的构型；
设置所述超低轨道卫星的工作轨道范围；
当所述超低轨道卫星的构型和所述工作轨道范围确定后，根据大气密度和来流方向确定所述大气阻力的值。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤一还包括：
根据所述超低轨道卫星的构型及所述超低轨道卫星的工作轨道范围，获取迎风面积和大气相对所述超低轨道卫星的速度；
根据所述工作轨道范围，获取所述工作轨道范围内的所述大气密度；
获取所述超低轨道卫星在工作轨道上的运行方向，并根据所述运行方向计算所述来流方向；
根据所述大气密度和所述来流方向计算阻力系数；
根据上述参数，估算所述大气阻力的值：
Figure PCTCN2019120878-appb-000001
其中，F为所述大气阻力的值，C d为所述阻力系数；ρ为所述工作轨道范围内的大气密度；S为所述迎风面积；v为所述大气相对卫星的速度；
根据所述大气阻力的值，得到所述大气阻力的量级。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤一还包括：
计算所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的变化趋势，根据所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的所述变化趋势，估算所述超低轨道卫星超低轨运行的一个轨道周期内的平均大气阻力。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤二包括：
所述惯性加速度测量系统的噪声包括航天器受大气阻力加速度扰动，以及多项其他扰动；
对惯性加速度测量系统的噪声进行分解，得出所述航天器受大气阻力加速度扰动；
根据所述航天器受大气阻力加速度扰动，评估所述航天器受大气阻力加速度扰动在所述惯性加速度测量系统的噪声中的比重，以使各项所述其他扰动的量级比所述航天器受大气阻力加速度扰动的量级低一级。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤二还包括：
获取所述超低轨道卫星的质量，根据所述大气阻力的值与所述超低轨道卫星的质量计算得到航天器受气动阻力加速度扰动：
σ1＝F/m，
其中，F为所述大气阻力的值，m为所述超低轨道卫星的质量。
可选的，在所述的超低轨道卫星轨道自主维持方法中，其中
所述其他扰动包括下列各项至少之一：所述惯性传感器的残余加速度 噪声、测试质量直接加速度扰动、航天器与测试质量耦合刚度误差、以及高频噪声与量化噪声之和；以及
其中所述步骤二还包括：获取所述惯性加速度测量系统的精度，根据所述惯性加速度测量系统的精度获取下列各项至少之一：所述惯性传感器的残余加速度噪声、所述测试质量直接加速度扰动、所述航天器与测试质量耦合刚度误差、以及所述高频噪声与量化噪声之和。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤三包括：
设置小推力执行系统的参数，其中所述小推力执行系统的参数包括下列各项至少之一：小推力执行系统的控制频率、推力器连续工作时间、以及小推力执行系统的推力器设计裕度，根据所述控制频率及所述推力器连续工作时间确定所述小推力执行系统的推力器设计裕度；
根据所述惯性加速度测量系统噪声分析结果，设置所述小推力执行系统的推力：
f＝K*σ；
其中：f为所述推力，K为所述推力器设计裕度，σ为所述惯性加速度测量系统的噪声。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤三还包括：
对所述惯性加速度测量系统进行在轨标定，得到标定后的惯性加速度输出结果包括：对所述惯性传感器进行标定，设置所述超低轨道卫星的姿态，所述惯性传感器多次测量所述超低轨道卫星的加速度，得到多个第一加速度测量值，计算所述多个第一加速度测量值的平均值，所述平均值作为所述标定后的惯性加速度输出结果；
对所述小推力执行系统进行在轨标定包括：完成所述惯性加速度测量系统在轨标定后，对所述推力器进行标定，所述推力器单次点火，所述惯性传感器测量所述超低轨道卫星的加速度，得到第二加速度测量值，根据所述超低轨道卫星的质量和所述多个第二加速度测量值，计算所述推力器 输出结果的各个控制周期的状态量。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤三包括：
对所述惯性加速度测量系统进行在轨标定时，所述惯性传感器测量所述超低轨道卫星的加速度的次数大于100次；
对所述惯性加速度测量系统进行在轨标定前，将所述超低轨道卫星姿态调整至指标要求精度以内，以减小标定时间内所述超低轨道卫星姿态变化，以减少所述超低轨道卫星姿态变化引起的外界干扰在敏感轴方向的投影分量；
对所述惯性加速度测量系统进行在轨标定时，设置所述超低轨道卫星处于姿态自由漂移模式，以减少所述超低轨道卫星姿态控制对惯性传感器测量标定引起的耦合干扰；
根据所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的变化趋势和所述超低轨道卫星超低轨运行的一个轨道周期内的平均大气阻力对所述小推力执行系统进行在轨标定。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤四包括：
根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法，计算当前时刻小推力执行系统的推力器输出结果，进行闭环控制：
X K+1＝A ob*X k+B ob*F k+L ob*Acc K+1*m
其中，X k为所述推力器输出结果的第k个控制周期的状态量；
A ob为推力器安装矩阵；
L ob为所述惯性加速度测量系统的安装矩阵；
B ob为无拖曳力矩输出矩阵；
F k为第k个控制周期施加的无拖曳控制推力；
Acc K+1为在第k+1个控制周期内由惯性传感器测量得到的预处理后的相对加速度；
m为卫星质量。
可选的，在所述的超低轨道卫星轨道自主维持方法中，所述超低轨道卫星轨道自主维持方法还包括步骤五，所述步骤五包括：
根据所述惯性传感器测量得到的预处理后的相对加速度，计算出气动力矩为主导的超低轨环境下的非保守力。
在本发明提供的超低轨道卫星轨道自主维持方法中，通过设置卫星的工作轨道范围，估算大气阻力的量级，分析惯性加速度测量系统的噪声的量级，根据惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对惯性加速度测量系统和小推力执行系统进行在轨标定，根据标定后的惯性加速度输出结果，设置小推力执行系统的轨控小推力输出算法，实现了一种基于持续小推力的卫星轨道自主维持方法，满足了卫星对超低轨道长期高精度维持需求，满足了运行在200～300km超低轨道卫星的长期轨道自主维持需求。
具体的，通过根据卫星的轨控策略确定小推力执行系统的控制频率及小推力执行系统的推力器连续工作时间，根据控制频率及推力器连续工作时间确定小推力执行系统的推力器设计裕度，根据惯性加速度测量系统噪声分析结果，设置小推力执行系统的推力，以及对惯性加速度测量系统进行在轨标定，实现了根据惯性传感器提供的相对加速度，反演出气动力矩为主导的超低轨环境下的非保守力，设计了基于持续小推力的轨道维持算法，轨道维持精度在200m量级，满足了超低轨道卫星对轨道长期自主维持的需求，并提高了轨道维持精度。
本发明的超低轨道卫星轨道自主维持方法具有较强的理论应用和工程实用价值，效果明显，由于主要是对软件进行改进，无需更换硬件，便于工程实现。
附图说明
图1是本发明一实施例的超低轨道卫星轨道自主维持方法流程示意图。
具体实施方式
以下结合附图和具体实施例对本发明提出的超低轨道卫星轨道自主维持方法作进一步详细说明。根据下面说明和权利要求书，本发明的优点和特征将更清楚。需说明的是，附图均采用非常简化的形式且均使用非精准的比例，仅用以方便、明晰地辅助说明本发明实施例的目的。
本发明的核心思想在于提供一种超低轨道卫星轨道自主维持方法，以解决现有的超低轨卫星的轨道维持精度较低的问题。
为实现上述思想，本发明提供了一种超低轨道卫星轨道自主维持方法，所述超低轨道卫星轨道自主维持方法包括：步骤一、设置卫星的工作轨道范围，估算大气阻力的量级；步骤二：根据所述大气阻力的量级，分析惯性加速度测量系统的噪声的量级，得到惯性加速度测量系统噪声分析结果；步骤三：根据所述惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对所述惯性加速度测量系统和所述小推力执行系统进行在轨标定，得到标定后的惯性加速度输出结果；以及步骤四：根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法。
<实施例一>
本实施例提供一种超低轨道卫星轨道自主维持方法，如图1所示，所述超低轨道卫星轨道自主维持方法包括：步骤一、设置超低轨道卫星的工作轨道范围，估算大气阻力的量级；步骤二：根据所述大气阻力的量级，分析惯性加速度测量系统的噪声的量级，得到惯性加速度测量系统噪声分析结果；步骤三：根据所述惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对所述惯性加速度测量系统和所述小推力执行系统进行在轨标定，得到标定后的惯性加速度输出结果；以及步骤四：根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法。
具体的，在所述的超低轨道卫星轨道自主维持方法中，所述惯性加速度测量系统包括惯性传感器，所述惯性传感器用于测量所述超低轨道卫星的加速度；和/或所述小推力执行系统包括推力器，所述推力器用于提供所述超低轨道卫星的动力。在所述的超低轨道卫星轨道自主维持方法中，所述步骤一包括：获取所述超低轨道卫星的构型；设置所述超低轨道卫星的 工作轨道范围；当所述超低轨道卫星的构型和所述工作轨道范围确定后，根据大气密度和来流方向确定所述大气阻力的值。在所述的超低轨道卫星轨道自主维持方法中，所述步骤一还包括：根据所述超低轨道卫星的构型及所述超低轨道卫星的工作轨道范围，获取迎风面积和大气相对所述超低轨道卫星的速度；根据所述工作轨道范围，获取所述工作轨道范围内的所述大气密度；获取所述超低轨道卫星在工作轨道上的运行方向，并根据所述运行方向计算所述来流方向；根据所述大气密度和所述来流方向计算阻力系数；根据上述参数，估算所述大气阻力的值：
Figure PCTCN2019120878-appb-000002
其中，F为所述大气阻力的值，C d为所述阻力系数；ρ为所述工作轨道范围内的大气密度；S为所述迎风面积；v为所述大气相对卫星的速度；根据所述大气阻力的值，得到所述大气阻力的量级。在所述的超低轨道卫星轨道自主维持方法中，所述步骤一还包括：计算所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的变化趋势，根据所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的所述变化趋势，估算所述超低轨道卫星超低轨运行的一个轨道周期内的平均大气阻力。
进一步的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤二包括：所述惯性加速度测量系统的噪声包括航天器受大气阻力加速度扰动，以及多项其他扰动；对惯性加速度测量系统的噪声进行分解，得出所述航天器受大气阻力加速度扰动；根据所述航天器受大气阻力加速度扰动，评估所述航天器受大气阻力加速度扰动在所述惯性加速度测量系统的噪声中的比重，以使各项所述其他扰动的量级比所述航天器受大气阻力加速度扰动的量级低一级。在所述的超低轨道卫星轨道自主维持方法中，所述步骤二还包括：获取所述超低轨道卫星的质量，根据所述大气阻力的值与所述超低轨道卫星的质量计算得到航天器受气动阻力加速度扰动：
σ1＝F/m，
其中，F为所述大气阻力的值，m为所述超低轨道卫星的质量。在所述的超低轨道卫星轨道自主维持方法中，其中：所述其他扰动包括下列各 项至少之一：所述惯性传感器的残余加速度噪声、测试质量直接加速度扰动、航天器与测试质量耦合刚度误差、以及高频噪声与量化噪声之和；以及所述步骤二还包括获取所述惯性加速度测量系统的精度，根据所述惯性加速度测量系统的精度获取下列各项至少之一：所述惯性传感器的残余加速度噪声、所述测试质量直接加速度扰动、所述航天器与测试质量耦合刚度误差、以及所述高频噪声与量化噪声之和。
更进一步的，在所述的超低轨道卫星轨道自主维持方法中，所述步骤三包括：设置小推力执行系统的参数，其中所述小推力执行系统的参数包括下列各项至少之一：小推力执行系统的控制频率、推力器连续工作时间、以及小推力执行系统的推力器设计裕度，根据所述控制频率及所述推力器连续工作时间确定所述小推力执行系统的推力器设计裕度；根据所述惯性加速度测量系统噪声分析结果，设置所述小推力执行系统的推力：
f＝K*σ；
其中：f为所述推力，K为所述推力器设计裕度，σ为所述惯性加速度测量系统的噪声。在所述的超低轨道卫星轨道自主维持方法中，所述步骤三还包括：对所述惯性加速度测量系统进行在轨标定，得到标定后的惯性加速度输出结果包括：对所述惯性传感器进行标定，设置所述超低轨道卫星的姿态，所述惯性传感器多次测量所述超低轨道卫星的加速度，得到多个第一加速度测量值，计算所述多个第一加速度测量值的平均值，所述平均值作为所述标定后的惯性加速度输出结果；对所述小推力执行系统进行在轨标定包括：完成所述惯性加速度测量系统在轨标定后，对所述推力器进行标定，所述推力器单次点火，所述惯性传感器测量所述超低轨道卫星的加速度，得到第二加速度测量值，根据所述超低轨道卫星的质量和所述多个第二加速度测量值，计算所述推力器输出结果的各个控制周期的状态量。在所述的超低轨道卫星轨道自主维持方法中，所述步骤三包括：对所述惯性加速度测量系统进行在轨标定时，所述惯性传感器测量所述超低轨道卫星的加速度的次数大于100次；对所述惯性加速度测量系统进行在轨标定前，将所述超低轨道卫星姿态调整至指标要求精度以内，以减小标定 时间内所述超低轨道卫星姿态变化，以减少所述超低轨道卫星姿态变化引起的外界干扰在敏感轴方向的投影分量；对所述惯性加速度测量系统进行在轨标定时，设置所述超低轨道卫星处于姿态自由漂移模式，以减少所述超低轨道卫星姿态控制对惯性传感器测量标定引起的耦合干扰；根据所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的变化趋势和所述超低轨道卫星超低轨运行的一个轨道周期内的平均大气阻力对所述小推力执行系统进行在轨标定。
另外，在所述的超低轨道卫星轨道自主维持方法中，所述步骤四包括：根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法，计算当前时刻小推力执行系统的推力器输出结果，进行闭环控制：
X K+1＝A ob*X k+B ob*F k+L ob*Acc K+1*m
其中，X k为所述推力器输出结果的第k个控制周期的状态量；A ob为推力器安装矩阵；L ob为所述惯性加速度测量系统的安装矩阵；B ob为无拖曳力矩输出矩阵；F k为第k个控制周期施加的无拖曳控制推力；Acc K+1为在第k+1个控制周期内由惯性传感器测量得到的预处理后的相对加速度；m为卫星质量。
最后，在所述的超低轨道卫星轨道自主维持方法中，所述超低轨道卫星轨道自主维持方法还包括步骤五，所述步骤五包括：根据所述惯性传感器测量得到的预处理后的相对加速度，计算出气动力矩为主导的超低轨环境下的非保守力。
在本发明提供的超低轨道卫星轨道自主维持方法中，通过设置卫星的工作轨道范围，估算大气阻力的量级，分析惯性加速度测量系统的噪声的量级，根据惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对惯性加速度测量系统和小推力执行系统进行在轨标定，根据标定后的惯性加速度输出结果，设置小推力执行系统的轨控小推力输出算法，实现了一种基于持续小推力的卫星轨道自主维持方法，满足了卫星对超低轨道长期高精度维持需求，满足了运行在200～300km超低轨道卫星的长期轨道自主维持需求。
具体的，通过根据卫星的轨控策略确定小推力执行系统的控制频率及小推力执行系统的推力器连续工作时间，根据控制频率及推力器连续工作时间确定小推力执行系统的推力器设计裕度，根据惯性加速度测量系统噪声分析结果，设置小推力执行系统的推力，以及对惯性加速度测量系统进行在轨标定，实现了根据惯性传感器提供的相对加速度，反演出气动力矩为主导的超低轨环境下的非保守力，设计了基于持续小推力的轨道维持算法，轨道维持精度在200m量级，满足了超低轨道卫星对轨道长期自主维持的需求，并提高了轨道维持精度。
本发明的超低轨道卫星轨道自主维持方法具有较强的理论应用和工程实用价值，效果明显，由于主要是对软件进行改进，无需更换硬件，便于工程实现。
综上，上述实施例对超低轨道卫星轨道自主维持方法的不同构型进行了详细说明，当然，本发明包括但不局限于上述实施中所列举的构型，任何在上述实施例提供的构型基础上进行变换的内容，均属于本发明所保护的范围。本领域技术人员可以根据上述实施例的内容举一反三。
上述描述仅是对本发明较佳实施例的描述，并非对本发明范围的任何限定，本发明领域的普通技术人员根据上述揭示内容做的任何变更、修饰，均属于权利要求书的保护范围。

Claims (13)

  1. 一种超低轨道卫星轨道自主维持方法，其特征在于，所述超低轨道卫星轨道自主维持方法包括：
    步骤一、设置超低轨道卫星的工作轨道范围，估算大气阻力的量级；
    步骤二：根据所述大气阻力的量级，分析惯性加速度测量系统的噪声的量级，得到惯性加速度测量系统噪声分析结果；
    步骤三：根据所述惯性加速度测量系统噪声分析结果，设置小推力执行系统的参数，并对所述惯性加速度测量系统和所述小推力执行系统进行在轨标定，得到标定后的惯性加速度输出结果；以及
    步骤四：根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法。
  2. 如权利要求1所述的超低轨道卫星轨道自主维持方法，其特征在于，所述惯性加速度测量系统包括惯性传感器，所述惯性传感器用于测量所述超低轨道卫星的加速度；和/或
    所述小推力执行系统包括推力器，所述推力器用于提供所述超低轨道卫星的动力。
  3. 如权利要求2所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤一包括：
    获取所述超低轨道卫星的构型；
    设置所述超低轨道卫星的工作轨道范围；
    当所述超低轨道卫星的构型和所述工作轨道范围确定后，根据大气密度和来流方向确定所述大气阻力的值。
  4. 如权利要求3所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤一还包括：
    根据所述超低轨道卫星的构型及所述超低轨道卫星的工作轨道范围，获取迎风面积和大气相对所述超低轨道卫星的速度；
    根据所述工作轨道范围，获取所述工作轨道范围内的所述大气密度；
    获取所述超低轨道卫星在工作轨道上的运行方向，并根据所述运行方向计算所述来流方向；
    根据所述大气密度和所述来流方向计算阻力系数；
    根据上述参数，估算所述大气阻力的值：
    Figure PCTCN2019120878-appb-100001
    其中，F为所述大气阻力的值，C d为所述阻力系数；ρ为所述工作轨道范围内的大气密度；S为所述迎风面积；v为所述大气相对卫星的速度；
    根据所述大气阻力的值，得到所述大气阻力的量级。
  5. 如权利要求3所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤一还包括：
    计算所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的变化趋势，根据所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的所述变化趋势，估算所述超低轨道卫星超低轨运行的一个轨道周期内的平均大气阻力。
  6. 如权利要求2所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤二包括：
    所述惯性加速度测量系统的噪声包括航天器受大气阻力加速度扰动，以及多项其他扰动；
    对惯性加速度测量系统的噪声进行分解，得出所述航天器受大气阻力加速度扰动；
    根据所述航天器受大气阻力加速度扰动，评估所述航天器受大气阻力加速度扰动在所述惯性加速度测量系统的噪声中的比重，以使各项所述其他扰动的量级比所述航天器受大气阻力加速度扰动的量级低一级。
  7. 如权利要求6所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤二还包括：
    获取所述超低轨道卫星的质量，根据所述大气阻力的值与所述超低轨道卫星的质量计算得到航天器受气动阻力加速度扰动：
    σ1＝F/m，
    其中，F为所述大气阻力的值，m为所述超低轨道卫星的质量。
  8. 如权利要求6所述的超低轨道卫星轨道自主维持方法，其特征在 于：
    其中所述其他扰动包括下列各项至少之一：所述惯性传感器的残余加速度噪声、测试质量直接加速度扰动、航天器与测试质量耦合刚度误差、以及高频噪声与量化噪声之和；以及
    其中所述步骤二还包括：获取所述惯性加速度测量系统的精度，根据所述惯性加速度测量系统的精度获取下列各项至少之一：所述惯性传感器的残余加速度噪声、所述测试质量直接加速度扰动、所述航天器与测试质量耦合刚度误差、以及所述高频噪声与量化噪声之和。
  9. 如权利要求2所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤三包括：
    设置小推力执行系统的参数，其中所述小推力执行系统的参数包括下列各项至少之一：小推力执行系统的控制频率、推力器连续工作时间、以及小推力执行系统的推力器设计裕度，根据所述控制频率及所述推力器连续工作时间确定所述小推力执行系统的推力器设计裕度；
    根据所述惯性加速度测量系统噪声分析结果，设置所述小推力执行系统的推力：
    f＝K*σ；
    其中：f为所述推力，K为所述推力器设计裕度，σ为所述惯性加速度测量系统的噪声。
  10. 如权利要求5所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤三还包括：
    对所述惯性加速度测量系统进行在轨标定，得到标定后的惯性加速度输出结果包括：对所述惯性传感器进行标定，设置所述超低轨道卫星的姿态，所述惯性传感器多次测量所述超低轨道卫星的加速度，得到多个第一加速度测量值，计算所述多个第一加速度测量值的平均值，所述平均值作为所述标定后的惯性加速度输出结果；
    对所述小推力执行系统进行在轨标定包括：完成所述惯性加速度测量系统在轨标定后，对所述推力器进行标定，所述推力器单次点火，所述惯 性传感器测量所述超低轨道卫星的加速度，得到第二加速度测量值，根据所述超低轨道卫星的质量和所述多个第二加速度测量值，计算所述推力器输出结果的各个控制周期的状态量。
  11. 如权利要求10所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤三包括：
    对所述惯性加速度测量系统进行在轨标定时，所述惯性传感器测量所述超低轨道卫星的加速度的次数大于100次；
    对所述惯性加速度测量系统进行在轨标定前，将所述超低轨道卫星姿态调整至指标要求精度以内，以减小标定时间内所述超低轨道卫星姿态变化，以减少所述超低轨道卫星姿态变化引起的外界干扰在敏感轴方向的投影分量；
    对所述惯性加速度测量系统进行在轨标定时，设置所述超低轨道卫星处于姿态自由漂移模式，以减少所述超低轨道卫星姿态控制对惯性传感器测量标定引起的耦合干扰；
    根据所述来流方向在所述超低轨道卫星超低轨运行的一个轨道周期内的变化趋势和所述超低轨道卫星超低轨运行的一个轨道周期内的平均大气阻力对所述小推力执行系统进行在轨标定。
  12. 如权利要求11所述的超低轨道卫星轨道自主维持方法，其特征在于，所述步骤四包括：
    根据所述标定后的惯性加速度输出结果，设置所述小推力执行系统的轨控小推力输出算法，计算当前时刻小推力执行系统的推力器输出结果，进行闭环控制：
    X K+1＝A ob*X k+B ob*F k+L ob*Acc K+1*m
    其中，X k为所述推力器输出结果的第k个控制周期的状态量；
    A ob为推力器安装矩阵；
    L ob为所述惯性加速度测量系统的安装矩阵；
    B ob为无拖曳力矩输出矩阵；
    F k为第k个控制周期施加的无拖曳控制推力；
    Acc K+1为在第k+1个控制周期内由惯性传感器测量得到的预处理后的 相对加速度；
    m为卫星质量。
  13. 如权利要求12所述的超低轨道卫星轨道自主维持方法，其特征在于，所述超低轨道卫星轨道自主维持方法还包括步骤五，所述步骤五包括：
    根据所述惯性传感器测量得到的预处理后的相对加速度，计算出气动力矩为主导的超低轨环境下的非保守力。
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CN114771873B (zh) * 2022-03-24 2024-05-03 北京控制工程研究所 一种超低轨卫星轨道自主精确维持方法
CN115675919A (zh) * 2022-10-31 2023-02-03 北京控制工程研究所 一种用于卫星主动指向超静平台的在轨标定方法
CN115675919B (zh) * 2022-10-31 2024-05-31 北京控制工程研究所 一种用于卫星主动指向超静平台的在轨标定方法

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