含固定转速驱动电机的挠性卫星自旋稳定性分析

分析一类含驱动电机挠性卫星的自旋稳定性,应用不变集定理证明了系统轨迹收敛于系统平衡点集,并通过构造正定的单调递减的类能量函数来判断各平衡点的稳定性质。获得的结论包括：由于挠性振动阻尼特性,在无外扰力矩作用时,含固定转速驱动电机的挠性卫星存在稳定的自旋;不含驱动电机的挠性卫星,稳定的自旋轴为其最大主惯量轴;驱动电机可改变挠性卫星的自旋轴,给出了稳定的自旋轴与电机转速方向夹角公式,该夹角为锐角;当电机转动角动量幅值远小于系统角动量时,稳定自旋轴接近最大主惯量轴;电机转动角动量幅值越接近系统角动量,稳定自旋轴越接近电机转动方向。文中也给出了稳定自旋轴平行电机转轴的条件。     

航天器环绕小行星Ivar的运动分析

将小行星Ivar近似为三轴椭球体，给出了非球形引力势函数，建立了航天器环绕小行星Ivar的轨道动力学方程。利用Jacobi积分常数绘制了航天器在Ivar周围的零速度曲线，并分析了航天器的可能运动区域，给出了航天器不碰撞小行星Ivar的边界条件及不同偏心率下的近拱点半径。分析了小行星Ivar扁率和椭率对环绕轨道的影响，数学仿真结果表明：在一个轨道周期内，顺行轨道的开普勒能量、轨道角动量、偏心率和近拱点半径变化较大，而逆行轨道的相应参数变化较小。     

不同聚集因素对燃烧室凝相颗粒粒度分布的影响

设计了一种凝相颗粒收集装置,可对燃烧室内不同聚集状态下的粒子进行完整的收集,并对颗粒的形态影响较小。针对HTPB推进剂,开展了不同流通面积和颗粒浓度条件下的粒子收集实验,利用扫描电镜和激光粒度分析仪,对收集到的粒子进行了分析。结果表明,不同聚集状态下的凝相颗粒粒径均分布在0.1～200μm之间,随着流通面积的减小,小于3μm的颗粒逐渐减少,3～10μm之间的颗粒变化不大,大于10μm的颗粒逐渐增加,颗粒平均粒径d43和d50均逐渐增大,且d43和d50的变化率也逐渐增加;随着颗粒浓度的增加,颗粒平均粒径d43和d50增大较明显。     

Optimal thrust program for deviating an Earth-pointing asteroid

The purpose of this study is to investigate the optimal thrust program for deviating an asteroid when it is flying directly toward or crossing the Earth. Under some assumptions, the problem can be considered as a two-body problem of Earth-asteroid system. The initial relative speed and distance are specified to be 10 km/s and 150 times of Earth's radius, respectively. We have about 1 day (or, 93681 seconds exactly) to take action. If a single impulse is applied to the asteroid at the specified initial point, the required impulse to obtain a miss distance of 2 times of Earth's radius is 169.5 m/s per kg mass. For an asteroid of 10 m in diameter, the total impulse required is 3.02 × 10⁸ m/s. It needs a typical large launching rocket to provide the total impulse. When the asteroid is larger or the initial distance is shorter, the number of launching rockets required increases rapidly. For further analysis with physical and engineering constraints imposed, we shall have to use the variational formulation method.     

Optimal nodal flyby with near-Earth asteroids using electric sail

The aim of this paper is to quantify the performance of an Electric Solar Wind Sail for accomplishing flyby missions toward one of the two orbital nodes of a near-Earth asteroid. Assuming a simplified, two-dimensional mission scenario, a preliminary mission analysis has been conducted involving the whole known population of those asteroids at the beginning of the 2013 year. The analysis of each mission scenario has been performed within an optimal framework, by calculating the minimum-time trajectory required to reach each orbital node of the target asteroid. A considerable amount of simulation data have been collected, using the spacecraft characteristic acceleration as a parameter to quantify the Electric Solar Wind Sail propulsive performance. The minimum time trajectory exhibits a different structure, which may or may not include a solar wind assist maneuver, depending both on the Sun-node distance and the value of the spacecraft characteristic acceleration. Simulations show that over 60% of near-Earth asteroids can be reached with a total mission time less than 100 days, whereas the entire population can be reached in less than 10 months with a spacecraft characteristic acceleration of 1 mm/s².     

L׳ADROIT – A spaceborne ultraviolet laser system for space debris clearing

Small (1–10 cm) debris in low Earth orbit (LEO) are extremely dangerous, because they spread the breakup cascade. Pulsed laser active debris removal using laser ablation jets on target is the most cost-effective way to re-enter the small debris. No other solutions address the whole problem of large (~100 cm, 1 t) as well as small debris. Physical removal of small debris (by nets, tethers and so on) is uneconomical because of the energy cost of matching orbits. In this paper, we present a completely new proposal relative to our earlier work. This new approach uses rapid, head-on interaction in 10–40 s rather than 4 minutes, using 20–40 kW bursts of 100 ps, 355 nm UV pulses from a 1.5 m diameter aperture on a space-based station in LEO. The station employs “heat-capacity” laser mode with low duty cycle to create an adaptable, robust, dual-mode system which can lower or raise large derelict objects into less dangerous orbits, as well as clear out the small debris in a 400-km thick LEO band. Time-average laser optical power is less than 15 kW. The combination of short pulses and UV wavelength gives lower required fluence on target as well as higher momentum coupling coefficient. An orbiting system can have short range because of high interaction rate deriving from its velocity through the debris field. This leads to much smaller mirrors and lower average power than the ground-based systems we have considered previously. Our system also permits strong defense of specific assets. Analysis gives an estimated cost less than $1 k each to re-enter most small debris in a few months, and about 280 k$ each to raise or lower 1-ton objects by 40 km. We believe it can do this for 2000 such large objects in about four years. Laser ablation is one of the few interactions in nature that propel a distant object without any significant reaction on the source.     

Deriving the spin rate/orientation from the quiescent spacecraft ABRIXAS using optical observations

With the loss of the battery system of ABRIXAS shortly after launch and consequent absence of telemetry, there was urgent need to determine ABRIXAS' spin rate and orientation in order to assess the possibility of re-establishing telemetry during periods of full-Sun orbits. We, therefore, conducted optical and video observations of ABRIXAS passages with a 1-second time resolution, and later simulated the optical appearance of ABRIXAS for several passages based on a three-dimensional model of the reflectivity properties of the 2.5m × 1.8m × 1.2m size-satellite. Here we present the results of the comparison of our grid of light curve models with the observations and show (i) how the spin rate of ABRIXAS slowed down between June and December 1999 and (ii) what information can be deduced on the temporal change of the orientation of the spin axis. We conclude with discussing the benefit of using ground based optical observation as a cost effective way to develop information about the orientation of a satellite when there is no telemetry.     

Photonic spin control for solar wind electric sail

The electric solar wind sail (E-sail) is a novel, efficient propellantless propulsion concept which utilises the natural solar wind for spacecraft propulsion with the help of long centrifugally stretched charged tethers. The E-sail requires auxiliary propulsion applied to the tips of the main tethers for creating the initial angular momentum and possibly for modifying the spinrate later during flight to counteract the orbital Coriolis effect and possibly for mission specific reasons. We introduce the possibility of implementing the required auxiliary propulsion by small photonic blades (small radiation pressure solar sails). The blades would be stretched centrifugally. We look into two concepts, one with and one without auxiliary tethers. The use of small photonic sails has the benefit of providing sufficient spin modification capability for any E-sail mission while keeping the technology fully propellantless. We conclude that small photonic sails appear to be a feasible and attractive solution to E-sail spinrate control.     

Rosetta visits asteroid (21)Lutetia

The International Rosetta Mission, cornerstone of the European Space Agency Scientific Programme, was launched on 2nd March 2004 to its 10 years journey to comet Churyumov–Gerasimenko. Rosetta will reach the comet in summer 2014, orbit it for about 1.5 years down to distances of a few Kilometres and deliver the Lander Philae onto its surface. After its successful asteroid fly-by in September 2008, Rosetta came back to Earth, for the final gravity acceleration towards its longest heliocentric orbit, up to a distance of 5.3 AU. It is during this phase that Rosetta crossed for the second time the main asteroids belt and performed a close encounter with asteroid (21)Lutetia on the 10th of July 2010 at a distance of ca. 3160 km and a relative velocity of 15 km/s. The payload complement of the spacecraft was activated to perform highly valuable scientific observations. The approach phase to the celestial body required a careful and accurate optical navigation campaign that will prove to be useful also for the comet approach phase. The experience gained with first asteroid flyby in 2008 was fed back into the operations definition and preparation for this highly critical phase; this concerns in particular the operations of the navigation camera for the close-loop autonomous asteroid tracking and of the main scientific camera for high resolution imaging. It was shortly after the flyby that Rosetta became the solar-powered spacecraft to have flown furthest from the Sun (>2.72 AU). This paper presents the activities carried out and planned for the definition, preparation and implementation of the asteroid flyby mission operations, including the test campaign conducted to improve the performance of the spacecraft and payload compared to the previous flyby. The results of the flyby itself are presented, with the operations implemented, the achieved performance and the lessons learned.     

Cassini Saturn-escape trajectories to Jupiter,Uranus, and Neptune

Potential encore-mission scenarios have been considered for the Cassini mission. In this paper we discuss one of the end-of-life scenarios in which the Cassini spacecraft could perform a Saturn escape via gravity assists from Titan. It is shown that such satellite-aided escape requires a small deterministic maneuver (e.g., Δv<50 m/s), but provides enough energy for the Cassini spacecraft to reach a range of targets in our Solar System, as close to the Sun as the asteroid belt or as far as the Kuiper belt. The escape sequence could be initiated from an arbitrary point during the on-going Cassini mission. Example tours are presented in which the final Titan flyby places the spacecraft into ballistic trajectories that reach Jupiter, Uranus, and Neptune. After years of heliocentric flight, the spacecraft could impact on the target gas giant or perform a flyby to escape from the Solar System (if not to another destination). The concept can be generalized to a new kind of missions, including nested-grand tours, which may involve satellite-aided captures and escapes at more than one planet.     

空间机器人抓取未知目标的质量特性参数辨识

根据动量守恒定理,研究了单臂自由飘浮空间机器人抓取未知目标的质量特性参数辨识问题。已有文献中基于动量守恒原理进行的参数辨识只适用于系统线动量和角动量为零的情况,当两者不为零时无法得到正确的辨识结果。为解决此问题,本文首先在惯性系而非本体系下推导得到了机械臂抓取未知目标后的系统线动量和角动量,保证了其各分量的守恒特性。然后基于此特性,通过在轨测量两个时刻机器人本体的线速度、角速度以及机械臂各关节的角速度和转角信息,建立了含有未知目标质量特性参数的动量增量方程。最后根据三组测量信息,从联立的动量增量方程组中求解得到未知目标的质量特性参数。数值仿真表明,此方法在系统线动量和角动量为零和不为零两种情况下都能实现高精度的参数辨识,同时还避免了方程求解的奇异问题。     

挠性卫星姿态的角动量反馈控制

利用动量交换装置的角动量测量和反馈来实现带有挠性附件卫星姿态的稳定控制是一个重要问题。为此，本文给出了角动量反馈控制器，指出了控制器参数的选取方法，分析了姿态控制误差与角动量测量误差的关系。结果表明，角动量反馈控制方法不仅能够保证姿态控制的稳定性，而且能够有效提高姿态控制的稳定性，并且对不确定挠性模态参数具有很强的鲁棒性。     

Conceptual design of a hypervelocity asteroid intercept vehicle (HAIV) and its flight validation mission

An intercept mission with nuclear explosives is the most effective of the practical mitigation options against the impact threat of near-Earth objects (NEOs) with a short warning time (e.g., much less than 10 years). The existing penetrated subsurface nuclear explosion technology limits the intercept velocity to less than approximately 300 m/s. Consequently, an innovative concept of blending a hypervelocity kinetic impactor with a subsurface nuclear explosion has been developed for optimal penetration, fragmentation, and dispersion of the target NEO. A proposed hypervelocity asteroid intercept vehicle (HAIV) consists of a kinetic-impact leader spacecraft and a follower spacecraft carrying nuclear explosives. This paper describes the conceptual development and design of a baseline HAIV system and its flight validation mission architecture for three mission cost classifications (e.g., $500 M, $1 B, and $1.5 B).     

细胞化卫星分布式角动量平衡力矩分配算法

针对细胞化卫星姿态控制系统中细胞模块数量多、控制器细胞通信压力大的问题,提出了一种分布式控制力矩分配算法。该力矩分配算法先将力矩分配问题转化为带耦合等式约束的凸优化问题,然后通过分布式的方式求解该优化问题。在力矩分配过程中,控制器细胞只需要提供期望力矩,因此降低了通信压力。同时,该算法在优化目标函数中引入角动量能力因子,使飞轮细胞可以根据自身的角动量水平决定其力矩输出。数值仿真结果表明,该分布式分配算法可以保证分配后的力矩与期望力矩相等,且实现了平衡飞轮系统角动量的目的,进而可以通过卸载部分细胞的角动量来卸载整个飞轮系统的角动量。     

Reducing mission operations costs through spacecraft autonomy: the near earth asteroid rendezvous (near) experience

With a maximum time of 12 days out of ground contact and a round-trip light time as high as 56 minutes, The Near Earth Asteroid Rendezvous (NEAR) spacecraft requires a moderate degree of onboard autonomy to react to faults and safe the spacecraft. Beyond the basic safing requirements, additional functions are carried out onboard. For example, on-board calculation of the Sun, Earth, asteroid, and spacecraft positions allow the spacecraft to autonomously orient itself for science and downlink operations. On-board autonomous momentum management during cruise relieves Mission Operations from planning, scheduling, and carrying out many manual momentum dumps. During development, additional operations, such as center-of-mass management during propulsive maneuvers and optical navigation were also considered for onboard autonomy on the NEAR spacecraft, but were not selected. The allocation of functions to onboard software or to ground operations involved tradeoffs such as development time for onboard software versus ground software, resource management, life cycle costs, and spacecraft safety.After two years of cruise operations, considerable experience with the NEAR autonomy system has accrued. The utility of some autonomous capabilities is greater than expected, others less so. Software uploads increased spacecraft autonomy in some cases, and the impact on Mission Operations can be assessed. Allocation of functions between spacecraft autonomy and ground operation during development of future missions can be improved by applying the lessons learned from the NEAR experience.     

自由浮动空间机械臂的最优控制

自由浮动空间机械臂所满足的动量矩守恒方程是一个非完整约束条件。首先把铰关节角速度作为输入，建立起系统的状态方程。然后把输入表示成待定系数的Ｆｏｕｒｉｅｒ 基组合，从而将最优控制转化为对代数问题的求解。最后用非线性最小二乘法迭代求出同时使载体和铰关节达到预定位形，而且使输入能量最小的最优控制。仿真算例说明，本文的方法是有效的。     

Dependence of density and burning rate of composite solid propellant on mixer size

This paper examines the effect of mixer size on the density of the propellant and it also explores the role of the propellant density with regards to burning rate and burn rate pressure index. Propellant samples were prepared in the four different size (15, 70, 200 and 1000 g) mixers with identical input of ingredients to examine the effect of mixer size on density and burning rate of the composite solid propellant. It was noticed that density and burning rate of the propellant changes significantly with change in the mixer size from 15 g to 1000 g. This is because for the smaller mixers the surface area to volume ratio is large and the actual percentage of aluminum and ammonium perchlorate (especially coarse) that goes into the propellant reduces. High burning rate with decrease in mixer size is also accompanied with increase in burn rate pressure index. The composition analysis was also carried out and it is noticed that the actual percentage of the ingredients is different from the intended percentage of the ingredients and percentage change is more when the mixer size decreases.     

一种使用磁力矩器管理双自旋航天器角动量的方法

针对一种新型采用双自旋设计进行对地遥感任务的航天器，提出了一种使用磁力矩器对旋转载荷角动量进行管理的方法。使用旋转载荷固连正交安装的三只磁力矩器，在旋转过程中连续输出周期性变化的磁矩，使自转一周过程中产生的合角动量方向仅沿自转轴方向，避免在其他两轴方向产生角动量累积。该方法在工程上可行性好，仿真结果表明旋转载荷角动量持续保持在期望值附近，不随时间累加。     

RF communications subsystem for the Radiation Belt Storm Probes mission

The NASA Radiation Belt Storm Probes (RBSP) mission, currently in Phase B, is a two-spacecraft, Earth-orbiting mission, which will launch in 2012. The spacecraft's S-band radio frequency (RF) telecommunications subsystem has three primary functions: provide spacecraft command capability, provide spacecraft telemetry and science data return, and provide accurate Doppler data for navigation. The primary communications link to the ground is via the Johns Hopkins University Applied Physics Laboratory's (JHU/APL) 18 m dish, with secondary links to the NASA 13 m Ground Network and the Tracking and Data Relay Spacecraft System (TDRSS) in single-access mode. The on-board RF subsystem features the APL-built coherent transceiver and in-house builds of a solid-state power amplifier and conical bifilar helix broad-beam antennas. The coherent transceiver provides coherency digitally, and controls the downlink data rate and encoding within its field-programmable gate array (FPGA). The transceiver also provides a critical command decoder (CCD) function, which is used to protect against box-level upsets in the C&DH subsystem. Because RBSP is a spin-stabilized mission, the antennas must be symmetric about the spin axis. Two broad-beam antennas point along both ends of the spin axis, providing communication coverage from boresight to 70°. An RF splitter excites both antennas; therefore, the mission is designed such that no communications are required close to 90° from the spin axis due to the interferometer effect from the two antennas. To maximize the total downlink volume from the spacecraft, the CCSDS File Delivery Protocol (CFDP) has been baselined for the RBSP mission. During real-time ground contacts with the APL ground station, downlinked files are checked for errors. Handshaking between flight and ground CFDP software results in requests to retransmit only the file fragments lost due to dropouts. This allows minimization of RF link margins, thereby maximizing data rate and thus data volume.     

航天器抓捕后复合体系统稳定的协调控制研究

针对航天器抓捕后由于系统质量特性和动量的改变而导致复合体系统失稳的问题,提出了两种基于角动量守恒的协调控制方法:关节阻尼控制和关节函数参数化协调控制。这两种方法通过对各关节和飞轮的速度进行协调规划和控制,实现对系统角动量的管理和重分配,在实现对目标进行停靠的同时,保证了基座的稳定性。两种方法各有优缺点,其中,关节函数参数化在实现系统稳定的同时还可使机械臂处于期望的臂型,以方便开展在轨维修等服务操作。所提出的方法将实际飞轮作为动量交换装置,具有很好的工程可实现性。仿真结果验证了方法的有效性。