导航卫星高精度太阳光压摄动分析建模及验证

为提高导航卫星精密定轨与轨道预报精度,提出了一种导航卫星太阳光压摄动的分析建模方法.相较于其他摄动因素模型完善且精度较高,光压摄动由于太阳活动导致太阳能量误差、卫星姿态控制误差和表面材料老化等问题,是最难以精确建模的摄动源,也是动力学模型最大的误差源.基于此,提出了一种基于卫星的姿态控制规律,通过分析法建立卫星太阳光压摄动模型,给出了光压摄动加速度在星体坐标系中的模型,并以GPSBlock IIR为例进行了验证.计算结果表明,该仿真分析法所建立的摄动模型与T30模型、ECOM模型精度接近,达到了光压建模研究的初步计算要求.     

Jason-1海洋观测卫星升空

美国东部时间 2 0 0 1年 12月 7日 ,在美国加利福尼亚州范登堡空军基地 ,法国空间中心和美国航宇局用美国波音公司的德尔他 - 2火箭 ,成功地发射了 Jason- 1海洋观测卫星 ,此次同时发射升空的还有美国的另一颗名为“Timed”的科学卫星。Jason- 1海洋观测卫星由法国空间中心和美国航宇局下属的喷气推进实验室联合研制。 Jason- 1卫星是“托佩克斯 /海神” (Topex/Poseidon)卫星的后继星 ,它将接替已经运行了 9年的 Topex/Poseidon卫星 ,为国际科学界迅速提供几个小时或几天内海洋状态的有关情况。Jason卫星计划是发展未来海洋观测和预测…     

磁定姿近地轨道卫星EKF与UKF算法比较

某近地轨道(LEO)卫星通过磁强计测量、飞轮控制和磁力矩器卸载,实现三轴姿态稳定.围绕该卫星姿态确定问题,对扩展卡尔曼滤波(EKF)和无迹卡尔曼滤波(UKF)算法进行综合分析对比,最后给出该卫星姿态确定实现方法.分析了地磁模型误差、磁强计测量误差及剩磁的影响,推导了四元数均值及协方差的计算方法.分别引入EKF和UKF进行姿态确定,并讨论了影响算法性能的因素.对两种方法的估计精度、收敛时间及计算需求等指标进行了对比.最后针对该卫星姿态确定要求和硬件条件,选用EKF算法实现姿态确定,满足对地指向精度优于1°,稳定度优于0.1(°)/s的设计指标.     

“嫦娥二号”卫星L2点角动量管理

对“嫦娥二号”(CE-2)卫星进入环绕日地拉格朗日L2点的李萨如轨道后卫星喷气卸载所产生的影响进行了研究.提出了一种利用光压力矩辅助卫星太阳电池翼角度调整进行角动量管理的方法.在轨试验表明,太阳光压强度足够对飞行在L2点环绕轨道上的CE-2卫星进行角动量管理,可以大幅度减少动量轮在轨喷气卸载的次数,有利于CE-2卫星的轨道维持.     

基于约化相对轨道拟平根数的长期稳定高精度卫星编队导航技术

针对基于星间相对测量的相对导航算法中由测量方程将相对轨道拟平根数转化为相对位置过程导致的模型非线性,提出一种基于约化相对轨道拟平根数的卫星编队导航方法.该方法通过编队卫星之间一段时间的切向漂移估计半长轴偏差,合理处理半长轴偏差对双星相对动力学的影响,克服了模型线性化造成的误差,能够实现长期稳定的高精度卫星编队导航.     

一种基于大力矩飞轮的敏捷卫星路径规划的姿态机动控制方法

遥感卫星对姿态机动能力的快速性提出了更高的需求,提出一种路径规划控制方法及在线计算模型,利用大力矩飞轮作为执行机构提供控制力矩,通过大量测试得到修正角度的模型,从而使卫星具备高精度敏捷姿态机动能力.通过数学仿真结果说明了本方法的有效性.     

高阶重力和力矩对空间太阳能电站运动的影响

相比于传统卫星,空间太阳能电站具有超大的尺寸,高阶重力和重力梯度对其轨道和姿态运动的影响将不能再忽略.文中以地球同步拉普拉斯轨道上的太阳塔式空间太阳能电站为例,研究了在考虑地球扁率的引力场中,高阶重力和力矩对空间太阳能电站姿轨运动的影响.首先将空间太阳能电站的重力势函数进行泰勒展开,并保留至四阶项;然后求出电站所受到的重力和重力梯度力矩,并给出其轨道和姿态运动方程;最后通过数值仿真来分析不同阶次的力和力矩对轨道运动的影响.结果表明:高阶力对卫星轨道的影响较大,可达到百米量级;高阶力矩对卫星姿态运动影响则较小,可忽略不计.     

转动载荷对卫星姿态的影响与控制研究

研究带有转动载荷和挠性附件的卫星姿态控制问题.基于具有广义坐标形式的牛顿 欧拉方法建立了卫星姿态动力学模型和转动载荷力矩模型,研究载荷产生的动不平衡力矩和静不平衡力矩的机理和特点.分析载荷干扰对卫星姿态的影响特性,给出基于传递函数进行拉氏变换以估算姿态抖动量的方法.分析卫星姿态控制系统设计干扰补偿控制器的条件,给出了控制器的工程设计方法.验证结果表明该方法有效且能提高卫星姿态精度.     

“嫦娥4号”中继星任务轨道确定问题初探

"嫦娥4号"任务将采用着陆器、巡视器和绕飞地月拉格朗日L2点中继星进行月球背面的探测,中继星已先期发射,进入环绕地月L2点的晕(Halo)轨道。在中继星使命轨道动力学模型的基础上,通过相关仿真工作,开展了中继星在Halo轨道上的摄动源量级及影响定轨预报的主要因素分析,结果表明:太阳光压摄动是其主要影响因素。为降低其相关影响,提高定轨精度,在太阳光压球模型的基础上,结合中继星在轨运行特点及其星体结构特点,提出了一种求解光压等效面积的方法。经仿真分析,使用修正后的太阳光压球模型进行定轨求解,速度精度可提升约一个量级。     

火星环绕器环火轨道的角动量管理方法

针对中国首次火星探测的任务需求和规划，结合环绕器质量特性和布局构型以及大椭圆轨道特性，分析了环绕器在环火飞行阶段所受的重力梯度力矩和光压力矩规律.空间干扰力矩累积导致飞轮转速上升，影响姿态机动能力，需要喷气卸载.为减少喷气卸载次数，在满足系统任务的基础上，设计了姿态偏置方案，利用光压力矩和重力梯度力矩相互抵消减少角动量累积.通过仿真表明该方法可以将空间干扰力矩累积量减少70%，节约卸载喷气的燃料消耗，延长航天器寿命.     

Drag and energy accommodation coefficients during sunspot maximum

Conditions appropriate to gas-surface interactions on satellite surfaces in orbit have not been successfully duplicated in the laboratory. However, measurements by pressure gauges and mass spectrometers in orbit have revealed enough of the basic physical chemistry that realistic theoretical models of the gas-surface interaction can now be used to calculate physical drag coefficients. The dependence of these drag coefficients on conditions in space can be inferred by comparing the physical drag coefficient of a satellite with a drag coefficient fitted to its observed orbital decay. This study takes advantage of recent data on spheres and attitude stabilized satellites to compare physical drag coefficients with the histories of the orbital decay of several satellites during the recent sunspot maximum. The orbital decay was obtained by fitting, in a least squares sense, the semi-major axis decay inferred from the historical two-line elements acquired by the US Space Surveillance Network. All the principal orbital perturbations were included, namely geopotential harmonics up to the 16th degree and order, third body attraction of the Moon and the Sun, direct solar radiation pressure (with eclipses), and aerodynamic drag, using the Jacchia-Bowman 2006 (JB2006) model to describe the atmospheric density. After adjusting for density model bias, a comparison of the fitted drag coefficient with the physical drag coefficient has yielded values for the energy accommodation coefficient as well as for the physical drag coefficient as a function of altitude during solar maximum conditions. The results are consistent with the altitude and solar cycle variation of atomic oxygen, which is known to be adsorbed on satellite surfaces, affecting both the energy accommodation and angular distribution of the reemitted molecules.     

倾斜轨道小卫星太阳高度角分析与机动方案设计

针对倾斜轨道小卫星必然会面对的光学姿态敏感器受阳光干扰的问题,分析太阳高度角的变化规律,提出了计算太阳高度角的经验公式,并在此基础上提出了平台偏航姿态机动方案,可以使光学姿态敏感器规避阳光干扰,并可以简化整星热控和帆板驱动设计,提高了整星的可靠性.     

Yarkovsky-Schach effect on space debris motion

The Yarkovsky-Schach effect is a small perturbation affecting Earth satellites and space debris illuminated by the Sun. It was first applied to the orbit of LAGEOS satellites as an explanation of the residuals in orbital elements. In this work, we carry out several numerical integration tests taking into consideration various orbit and rotation parameters, in order to analyse this effect in a broader context. The semi-major axis variations remain small and depend on the spin axis attitude with respect to the Sun. We show that the force amplitude is maximised for orbits inclined with $i\approx$?20–30°. We also observe the influence on other orbital elements, notably on the orbit inclination. However, these effects are clearly observed only on long timescales; in our simulations, we propagated the orbits for 200?y. The Yarkovsky-Schach effect is thus confirmed to have a minuscule magnitude. It should be taken into account in studies requiring high-precision orbit determination, or on expanded timescales.     

A deorbiter CubeSat for active orbital debris removal

This paper introduces a mission concept for active removal of orbital debris based on the utilization of the CubeSat form factor. The CubeSat is deployed from a carrier spacecraft, known as a mothership, and is equipped with orbital and attitude control actuators to attach to the target debris, stabilize its attitude, and subsequently move the debris to a lower orbit where atmospheric drag is high enough for the bodies to burn up. The mass and orbit altitude of debris objects that are within the realms of the CubeSat’s propulsion capabilities are identified. The attitude control schemes for the detumbling and deorbiting phases of the mission are specified. The objective of the deorbiting maneuver is to decrease the semi-major axis of the debris orbit, at the fastest rate, from its initial value to a final value of about 6471?km (i.e., 100?km above Earth considering a circular orbit) via a continuous low-thrust orbital transfer. Two case studies are investigated to verify the performance of the deorbiter CubeSat during the detumbling and deorbiting phases of the mission. The baseline target debris used in the study are the decommissioned KOMPSAT-1 satellite and the Pegasus rocket body. The results show that the deorbiting times for the target debris are reduced significantly, from several decades to one or two years.     

An empirical solar radiation pressure model for satellites moving in the orbit-normal mode

We present a family of empirical solar radiation pressure (SRP) models suited for satellites orbiting the Earth in the orbit normal (ON) mode. The proposed ECOM-TB model describes the SRP accelerations in the so-called terminator coordinate system. The choice of the coordinate system and the SRP parametrization is based on theoretical assumptions and on simulation results with a QZS-1-like box-wing model, where the SRP accelerations acting on the solar panels and on the box are assessed separately. The new SRP model takes into account that in ON-mode the incident angle of the solar radiation on the solar panels is not constant like in the yaw-steering (YS) attitude mode. It depends on the elevation angle of the Sun above the satellite’s orbital plane. The resulting SRP vector acts, therefore, not only in the Sun-satellite direction, but has also a component normal to it. Both components are changing as a function of the incident angle. ECOM-TB has been used for precise orbit determination (POD) for QZS-1 and BeiDou2 (BDS2) satellites in medium (MEO) and inclined geosynchronous Earth orbits (IGSO) based on IGS MGEX data from 2014 and 2015. The resulting orbits have been validated with SLR, long-arc orbit fits, orbit misclosures, and by the satellite clock corrections based on the orbits. The validation results confirm that—compared to ECOM2—ECOM-TB significantly (factor 3–4) improves the POD of QZS-1 in ON-mode for orbits with different arc lengths (one, three, and five days). Moderate orbit improvements are achieved for BDS2 MEO satellites—especially if ECOM-TB is supported by pseudo-stochastic pulses (the model is then called ECOM-TBP). For BDS2 IGSOs, ECOM-TB with its 9 SRP parameters appears to be over-parameterized. For use with BDS2 IGSO spacecraft we therefore developed a minimized model version called ECOM-TBMP, which is based on the same axis decomposition as ECOM-TB, but has only 2 SRP parameters and is supported by pseudo-stochastic parameters, as well. This model shows a similar performance as ECOM-TB with short arcs, but an improved performance with (3-day) long-arcs. The new SRP models have been activated in CODE’s IGS MGEX solution in Summer 2018. Like the other ECOM models the ECOM-TB derivatives might be used together with an a priori model.     

Measuring atmospheric density using GPS–LEO tracking data

We present a method to estimate the total neutral atmospheric density from precise orbit determination of Low Earth Orbit (LEO) satellites. We derive the total atmospheric density by determining the drag force acting on the LEOs through centimeter-level reduced-dynamic precise orbit determination (POD) using onboard Global Positioning System (GPS) tracking data. The precision of the estimated drag accelerations is assessed using various metrics, including differences between estimated along-track accelerations from consecutive 30-h POD solutions which overlap by 6 h, comparison of the resulting accelerations with accelerometer measurements, and comparison against an existing atmospheric density model, DTM-2000. We apply the method to GPS tracking data from CHAMP, GRACE, SAC-C, Jason-2, TerraSAR-X and COSMIC satellites, spanning 12 years (2001–2012) and covering orbital heights from 400 km to 1300 km. Errors in the estimates, including those introduced by deficiencies in other modeled forces (such as solar radiation pressure and Earth radiation pressure), are evaluated and the signal and noise levels for each satellite are analyzed. The estimated density data from CHAMP, GRACE, SAC-C and TerraSAR-X are identified as having high signal and low noise levels. These data all have high correlations with anominal atmospheric density model and show common features in relative residuals with respect to the nominal model in related parameter space. On the contrary, the estimated density data from COSMIC and Jason-2 show errors larger than the actual signal at corresponding altitudes thus having little practical value for this study. The results demonstrate that this method is applicable to data from a variety of missions and can provide useful total neutral density measurements for atmospheric study up to altitude as high as 715 km, with precision and resolution between those derived from traditional special orbital perturbation analysis and those obtained from onboard accelerometers.     

挠性卫星轨控期间动力学与姿态控制

卫星轨道控制期间,轨道控制推力会激振挠性太阳帆板,从而也影响卫星的姿态。用拉格朗日方程建立了带有大型太阳帆板的卫星动力学模型,分析了卫星质心运动、姿态运动与挠性振动的耦合关系。根据姿态控制推力器的输出特性,设计了轨道控制期间卫星姿态控制方案。通过数学仿真验证了轨道控制推力对挠性帆板与卫星姿态的影响,验证了轨控期间姿态控制方案的有效性。     

基于光学测角数据的风云四号同步轨道卫星精密定轨

针对风云四号同步卫星的精密定轨和精度评估需求，首先利用地面光学测角数据对FY-4A卫星进行精密定轨，定轨后方位角和高度角的残差rms分别为0.25＂和0.45＂。与基于测距数据的轨道相比，位置精度在有测角数据的弧段内小于50m。进一步联合测角数据和测距数据对FY-4A卫星进行联合定轨，定轨后轨道重叠精度优于15m。利用联合定轨结果评估了基于测距数据的实时轨道产品精度，可以明显发现轨道精度随着测距数据的积累而逐步提高。     

一种空间协同探测系统的星座设计方法

研究了由多颗合成孔径雷达卫星和一颗可见光卫星构成的空间协同探测系统的工作模式.根据太阳同步轨道和冻结轨道的特点,结合近地轨道遥感卫星的应用需求,选择了轨道半长轴、偏心率、倾角、近心点幅角.考虑到卫星偏航控制对覆盖性能的影响,在不动的地心坐标系中推导了卫星观测方向与地表交点的表达式,提出了确定系统中各颗卫星的升交点赤经和过近心点时刻的算法.给出了包含两颗合成孔径雷达卫星与一颗可见光卫星的协同探测系统的星座设计结果,并利用国际公认的卫星软件工具包Satellite ToolKit进行了验证,表明该设计方法是正确的.     

Simulation of long-term rotational dynamics of large space debris: A TOPEX/Poseidon case study

Accurate knowledge of the rotational dynamics of a large space debris is crucial for space situational awareness (SSA), whether it be for accurate orbital predictions needed for satellite conjunction analyses or for the success of an eventual active debris removal mission charged with stabilization, capture and removal of debris from orbit. In this light, the attitude dynamics of an inoperative satellite of great interest to the space debris community, the joint French and American spacecraft TOPEX/Poseidon, is explored. A comparison of simulation results with observations obtained from high-frequency satellite range measurements is made, showing that the spacecraft is currently spinning about its minor principal axis in a stable manner. Predictions of the evolution of its attitude motion to 2030 are presented, emphasizing the uncertainty on those estimates due to internal energy dissipation, which could cause a change of its spin state in the future. The effect of solar radiation pressure and the eddy-current torque are investigated in detail, and insights into some of the satellite’s missing properties are provided. These results are obtained using a novel, open-source, coupled orbit-attitude propagation software, the Debris SPin/Orbit Simulation Environment (D-SPOSE), whose primary goal is the study of the long-term evolution of the attitude dynamics of large space debris.