一种提高空间实验室定轨精度的方法

以近地航天器轨道动力学为基础,建立变阻力系数大气摄动模型,设计了求解变阻力系数的算法。然后利用天宫一号飞行器的测轨数据进行计算,分析了空间实验室飞行高度的轨道特性,其中包括：大气模式密度误差、变阻力系数与空间环境关系、定轨残差和星历误差。在空间环境平静和磁暴的条件下,制定了多种求解变阻力系数的策略,解决了空间实验室长弧段定轨精度受限的问题,并在空间环境平静条件下实现了优于10米的定轨精度,在磁暴条件下实现了优于20米的定轨精度。     

Navigation support for the RadioAstron mission

A developed method of determination of orbital parameters allows one to estimate, along with orbit elements, some additional parameters that characterize solar radiation pressure and perturbing accelerations due to unloadings of reactiion wheels. A parameterized model of perturbing action of solar radiation pressure on the spacecraft motion is described (this model takes into account the shape, reflecting properties of surfaces, and spacecraft attitude). Some orbit determination results are presented obtained by the joint processing of radio measurements of slant range and Doppler, laser range measurements used to calibrate the radio measurements, optical observations of right ascension and declination, and telemetry data on spacecraft thrusters’ firings during an unloading of reaction wheels.     

Autonomous aerobraking for low-cost interplanetary missions

Aerobraking has previously been used to reduce the propellant required to deliver an orbiter to its desired final orbit. In principle, aerobraking should be possible around any target planet or moon having sufficient atmosphere to permit atmospheric drag to provide a portion of the mission ΔV, in lieu of supplying all of the required ΔV propulsively. The spacecraft is flown through the upper atmosphere of the target using multiple passes, ensuring that the dynamic pressure and thermal loads remain within the spacecraft's design parameters. NASA has successfully conducted aerobraking operations four times, once at Venus and three times at Mars. While aerobraking reduces the fuel required, it does so at the expense of time (typically 3–6 months), continuous Deep Space Network (DSN) coverage, and a large ground staff. These factors can result in aerobraking being a very expensive operational phase of the mission. However, aerobraking has matured to the point that much of the daily operation could potentially be performed autonomously onboard the spacecraft, thereby reducing the required ground support and attendant aerobraking related costs. To facilitate a lower-risk transition from ground processing to an autonomous capability, the NASA Engineering and Safety Center (NESC) has assembled a team of experts in aerobraking and interplanetary guidance and control to develop a high-fidelity, flight-like simulation. This simulation will be used to demonstrate the overall feasibility while exploring the potential for staff and DSN coverage reductions that autonomous aerobraking might provide. This paper reviews the various elements of autonomous aerobraking and presents an overview of the various models and algorithms that must be transformed from the current ground processing methodology to a flight-like environment. Additionally the high-fidelity flight software test bed, being developed from models used in a recent interplanetary mission, will be summarized.     

How we will go to Mars

This article studies the efficiency of ejecting waste generated by the life support system (LSS) of a manned spacecraft to reduce initial mass on low earth orbit. The spacecraft is used for a long-duration interplanetary mission and is equipped with either a chemical or a nuclear-thermal propulsion system. For this study we simulate an optimal control problem for a given spacecraft maneuver. An impulsive approximation of the optimal interplanetary spacecraft trajectory is assumed, which allows us to reduce the general optimal control problem to hierarchic structure of 'outer' and 'inner' subproblems. This structure is analyzed using the Pontryagin's Maximum principle. Numerical results, illustrating the efficiency of waste ejection are shown for typical Earth-Mars transfer trajectories. This results confirm in theory that using a waste ejection system makes an early manned Mars mission possible without having to design and build new, advanced biological LSS.     

Rosetta: The ESA comet rendezvous mission

Rosetta was selected in November 1993 for the ESA Cornerstone 3 mission, to be launched in 2003, dedicated to the exploration of the small bodies of the solar system (asteroids and comets). Following this selection, the Rosetta mission and its spacecraft have been completely reviewed: this paper presents the studies performed the proposed mission and the resulting spacecraft design.

Three mission opportunities have been identified in 2003–2004, allowing rendezvous with a comet. From a single Ariane 5 launch, the transfer to the comet orbit will be supported by planetary gravity assists (two from Earth, one from Venus or Mars); during the transfer sequence, two asteroid fly-bys will occur, allowing first mission science phases. The comet rendezvous will occur 8–9 years after launch; Rosetta will orbit around the comet and the main science mission phase will take place up to the comet perihelion (1–2 years duration).

The spacecraft design is driven (i) by the communication scenario with the Earth and its equipment, (ii) by the autonomy requirements for the long cruise phases which are not supported by the ground stations, (iii) by the solar cells solar array for the electrical power supply and (iv) by the navigation scenario and sensors for cruise, target approach and rendezvous phases. These requirements will be developed and the satellite design will be presented.     

木星探测器轨道计算中的动力学

针对木星探测任务中轨道计算所面临的轨道动力学问题进行研究。首先，对木星探测轨道计算中涉及的时空参考系进行概述，比较了木星动力学时(TDJ)与质心动力学时(TDB)的差异；其次，对木星的定向参数模型及转换关系进行描述，给出了木星天球参考系与固联参考系的转换关系；再次，对木星探测器定轨计算中所需考虑的动力学模型进行讨论，提供了各类摄动加速度计算所需的公式；最后，对木星探测器JUNO的星历数据进行动力学拟合，检验了动力学模型的正确性。     

一种用于实时轨道确定的NPF-SRCKF滤波算法

针对航天器实时轨道确定中建模复杂、计算量大、估计精度低的问题,设计了一种考虑地球J 2 摄动影响的基于NPF-SRCKF的实时轨道确定算法,该算法采用非线性预测滤波(NPF)对模型误差进行补偿修正,利用平方根容积卡尔曼滤波(SRCKF)算法对修正模型误差后的系统进行状态估计。针对单测站、双测站跟踪测量设计了不同的实时轨道确定算法。实验结果显示,将非线性预测滤波和平方根容积卡尔曼滤波结合在一起,运用于简化的实时轨道确定模型,能有效降低计算复杂度,改进数值运算的稳定性,提高状态估计的精度。     

Low cost planetary exploration: surrey lunar minisatellite and interplanetary platform missions

In order to meet the growing global requirement for affordable missions beyond Low Earth Orbit, two types of platform are under design at the Surrey Space Centre. The first platform is a derivative of Surrey's UoSAT-12 minisatellite, launched in April 1999 and operating successfully in-orbit. The minisatellite has been modified to accommodate a propulsion system capable of delivering up to 1700 m/s delta-V, enabling it to support a wide range of very low cost missions to LaGrange points, Near-Earth Objects, and the Moon. A mission to the Moon - dubbed “MoonShine” - is proposed as the first demonstration of the modified minisatellite beyond LEO. The second platform - Surrey's Interplanetary Platform - has been designed to support missions with delta-V requirements up to 3200 m/s, making it ideal for low cost missions to Mars and Venus, as well as Near Earth Objects (NEOs) and other interplanetary trajectories. Analysis has proved mission feasibility, identifying key challenges in both missions for developing cost-effective techniques for: spacecraft propulsion; navigation; autonomous operations; and a reliable safe mode strategy. To reduce mission risk, inherently failure resistant lunar and interplanetary trajectories are under study. In order to significantly reduce cost and increase reliability, both platforms can communicate with low-cost ground stations and exploit Surrey's experience in autonomous operations. The lunar minisatellite can provide up to 70 kg payload margin in lunar orbit for a total mission cost US$16–25 M. The interplanetary platform can deliver 20 kg of scientific payload to Mars or Venus orbit for a mission cost US$25–50 M. Together, the platforms will enable regular flight of payloads to the Moon and interplanetary space at unprecedented low cost. This paper outlines key systems engineering issues for the proposed Lunar Minisatellite and interplanetary Platform Missions, and describes the accommodation and performance offered to planetary payloads.     

Two-craft tether formation relative equilibria about circular orbits and libration points

The relative equilibria of a two spacecraft tether formation connected by line-of-sight elastic forces moving in the context of a restricted two-body system and a circularly restricted three-body system are investigated. For a two spacecraft formation moving in a central gravitational field, a common assumption is that the center of the circular orbit is located at the primary mass and the center of mass of the formation orbits around the primary in a great-circle orbit. The relative equilibrium is called great-circle if the center of mass of the formation moves on the plane with the center of the gravitational field residing on it; otherwise, it is called a nongreat-circle orbit. Previous research shows that nongreat-circle equilibria in low Earth orbits exhibit a deflection of about a degree from the great-circle equilibria when spacecraft with unequal masses are separated by 350 km. This paper studies these equilibria (radial, along-track and orbit-normal in circular Earth orbit and Earth–Moon Libration points) for a range of inter-craft distances and semi-major axes of the formation center of mass. In the context of a two-spacecraft Coulomb formation with separation distances on the order of dozens of meters, this paper shows that the equilibria deflections are negligible (less than 10^?6°) even for very heterogeneous mass distributions. Furthermore, the nongreat-circle equilibria conditions for a two spacecraft tether structure at the Lagrangian libration points are developed.     

The Use of a Solar Electrojet Propulsion System for Jupiter Satellite Injection

The problem of optimization of the interplanetary trajectory of flight for a multistage spacecraft with high- and low-thrust engines into the Jupiter satellite orbit is considered. Low-thrust engines (stationary plasma engines) are used on a heliocentric flight segment. Their operation is maintained with electric power supply from solar batteries. The principal feasibility of the realization of such a project is shown, and estimations of the mass of a spacecraft placed into Jupiter's satellite orbit are presented.     

Economics of telecommunications space segments

This paper proposes a complete model for assessing the economics of telecommunications satellite systems, accounting for spacecraft development and manufacturing, launch and operations in orbit. This allows to account for such parameters as the mass and lifetime of the satellites, the number and type of payloads, the number of satellites procured and launched, the spare policy, the launch vehicle, the insurances, the satellite average MTTF and the management of the space segment efforts.

The model is divided into four parts: the spacecraft mass model, the spacecraft procurement cost model, the MTTF model and the space segment cost-effectiveness model. It provides for the rapid solution of a number of problems within a wide range of parameters such as assessing the influence on space segment economics of —certain satellite technologies, —satellite and payload mass, —number of payloads per spacecraft, —satellite lifetime, or —spare policy.     

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.     

利用GPS非差观测值的GRACE卫星精密定轨

参照GPS精密单点定位（PPP，Precise Point Positioning）模型设计了一种新的卫星定轨方法一组合星载加速度计测量数据和IGS提供的GPS精密星历及精密钟差数据进行低轨卫星的精密定轨。利用星载加速度计提高卫星受力模型准确性，使动力法定轨精度和可靠性都得到提升。同时，采用多种改正技术提高GPS非差观测值测量精度，保证最终高精度卫星定轨。本文建立了卫星定轨的轨道滤波模型，得出了有益的结论，即采用星载加速度计测量卫星非保守力可提高卫星定轨精度，在ITRF2000参考系下三轴精度优于18cm。这种方法不需要在全球建市基准观测站．定轨设备简单．费用低廉．     

航天器碰撞分析方法研究

作为航天器碰撞与规避控制方法研究课题的一部分,为此分别建立了航天器及其它空间目标的轨道确定方法;并在此基础上,提出了可能与航天器碰撞的其它空间编目目标的初选准则和碰撞预报算法。实测数据计算表明,该方法对于航天器轨道设计以及飞行安全性分析等均是实用可行的。     

GPS-based relative navigation for the Proba-3 formation flying mission

The primary objective of the Proba-3 mission is to build a solar coronagraph composed of two satellites flying in close formation on a high elliptical orbit and tightly controlled at apogee. Both spacecraft will embark a low-cost GPS receiver, originally designed for low-Earth orbits, to support the mission operations and planning during the perigee passage, when the GPS constellation is visible. The paper demonstrates the possibility of extending the utilization range of the GPS-based navigation system to serve as sensor for formation acquisition and coarse formation keeping. The results presented in the paper aim at achieving an unprecedented degree of realism using a high-fidelity simulation environment with hardware-in-the-loop capabilities. A modified version of the flight-proven PRISMA navigation system, composed of two single-frequency Phoenix GPS receivers and an advanced real-time onboard navigation filter, has been retained for this analysis. For several-day long simulations, the GPS receivers are replaced by software emulation to accelerate the simulation process. Special attention has been paid to the receiver link budget and to the selection of a proper attitude profile. Overall the paper demonstrates that, despite a limited GPS tracking time, the onboard navigation filter gets enough measurements to perform a relative orbit determination accurate at the centimeter level at perigee. Afterwards, the orbit prediction performance depends mainly on the quality of the onboard modeling of the differential solar radiation pressure acting on the satellites. When not taken into account, this perturbation is responsible for relative navigation errors at apogee up to 50 m. The errors can be reduced to only 10 m if the navigation filter is able to model this disturbance with 70% fidelity.     

Mars Express spacecraft: design and development solutions for affordable planetary missions

The spacecraft designed to support the ESA Mars Express mission and its science payloads is customized around an existing avionics well suited to environmental and operational constraints of deep-space interplanetary missions. The reuse of the avionics initially developed for the Rosetta cometary program thanks to an adequate ESA cornerstone program budget paves the way for affordable planetary missions.The costs and schedule benefits inherited from reuse of up-to-date avionics solutions validated in the frame of other programs allows to focus design and development efforts of a new mission over the specific areas which requires customization, such as spacecraft configuration and payload resources. This design approach, combined with the implementation of innovative development and management solutions have enabled to provide the Mars Express mission with an highly capable spacecraft for a remarkably low cost. The different spacecraft subsystems are all based on adequate design solutions. The development plan ensures an exhaustive spacecraft verification in order to perform the mission at minimum risk. New management schemes contribute to maintain the mission within its limited funding.Experience and heritage gained on this program will allow industry to propose to Scientists and Agencies high performance, low-cost solutions for the ambitious Mars Exploration Program of the forthcoming decade.     

Motion and transport of solar cosmic rays in heliospheric traps: The event on January 28–31, 2001

The results of studying the enhancement of solar cosmic ray fluxes on January 28?C31, 2001 in a wide energy range are presented using the ACE spacecraft data. A comparative analysis of temporal variations of the fluxes of charged particles and of the interplanetary medium parameters (interplanetary magnetic field and solar wind) has been performed on the basis of the ??reflection?? model of motion, accumulation, and modulation of cosmic rays. It is shown that a magnetic trap for solar cosmic rays was created by a plasma stream and flare ejection from an active region in the western part of the solar disk. Particles of low energies (<10 MeV) were captured inside the trap; the dispersion of distribution of particles with different energies inside the trap being determined by its complicated magnetic structure. The power-low dependence of the time of maximum for the flux of particles on their energy is found, and softer energy spectrum inside the trap is explained.     

月球探测器LP精密定轨及月球重力场模型解算

对美国1998年发射的月球探测器LP任务阶段共19个月的双程测距测速轨道跟踪数据进行了精密定轨,对定轨结果通过轨道残差及重复弧段差异进行了精度评价。利用LP正常任务阶段三个月的轨道跟踪数据进行了月球重力场模型解算,通过重力场功率谱、轨道残差和月球自由空气重力异常对解算模型进行了精度评价。结果表明精密定轨及月球重力场模型解算合理。对进一步融合嫦娥一号轨道跟踪数据和LP数据解算自主的高精度月球重力场模型具有参考意义。
　
     

Rosetta enters hibernation

The International Rosetta Mission was launched on 2nd March 2004 on its 10 years journey to comet 67P/Churyumov–Gerasimenko. Rosetta will reach the comet in 2014, orbit it for about 1.5 years down to distances of a few kilometres and deliver the Lander Philae onto its surface.Following the fly-by of Asteroid (21-)Lutetia in 2010, Rosetta continued its travel towards the planned comet encounter in 2014. In this phase Rosetta became the solar-powered spacecraft that reached the largest Sun distances in history of spaceflight, up to an aphelion at 5.3 AU in October 2012. At distances above 4.5 AU the spacecraft's solar generator power is not sufficient to keep all spacecraft systems active. Therefore in June 2011 the spacecraft was spun up to provide gyroscopic stabilisation, and most of its on-board units, including those used for attitude control and communications, were switched off. Over this “hibernation” phase of about 2.5 years the spacecraft will keep a minimum of autonomy active to ensure maintenance of safe thermal conditions.After Lutetia fly-by, flight controllers had to tackle two anomalies that had significant impacts on the mission operations. A leak in the reaction control subsystem was confirmed and led to the re-definition of the operational strategy to perform the comet rendezvous manoeuvres planned for 2011 and 2014. Anomalous jumps detected in the estimated friction torque of two of the four reaction wheels used for attitude control forced the rapid adoption of measures to slow down the wheels degradation. This included in-flight re-lubrication activities and changes in the wheels operational speed regime.Once the troubleshooting of the two anomalies was completed, and the related operational scenarios were implemented, the first large (790 m/s) comet rendezvous manoeuvre was executed, split into several long burns in January and February 2011. The second burn was unexpectedly interrupted due to the anomalous behaviour of two thrusters, causing attitude off-pointing. Flight controllers modified the thrusters operation parameters in the on-board software and managed to re-start the sequence of burns and successfully complete the manoeuvre. After the manoeuvre, preparation for the critical spin-up and hibernation entry activities, planned for June 2011, began.This paper presents the activities carried out on Rosetta in the final year before hibernation entry. The major anomalies and the related troubleshooting and workaround solutions are detailed. Lessons learned from the operation of the first spacecraft operating with solar power at Jupiter-like distances from the Sun are presented and discussed.     

Relationship between the fluxes of relativistic electrons at geosynchronous orbit and the level of ULF activity on the Earth’s surface and in the solar wind during the 23rd solar activity cycle

We present the results of a cross-correlation analysis made on the basis of Spearman’s rank correlation method. The quantities to correlate are daily values of the fluence of energetic electrons at a geosynchronous orbit, intensities of ground and interplanetary ultra-low-frequency (ULF) oscillations in the Pc5 range, and parameters of the solar wind. The period under analysis is the 23rd cycle of solar activity, 1996–2006. Daily (from 6 h to 18 h of LT) magnetic data at two diametrically opposite observatories of the Intermagnet network are taken as ground-based measurements. The fluxes of electrons with energies higher than 2 MeV were measured by the geosynchronous GOES satellites. The data of magnetometers and plasma instruments installed on ACE and WIND spacecraft were used for analysis of the solar wind parameters and of the oscillations of the interplanetary magnetic field (IMF). Some results elucidating the role played by interplanetary ULF waves in the processes of generation of magneospheric oscillations and acceleration of energetic electrons are obtained. Among them are (i) high and stable correlation of ground ULF oscillations with waves in the solar wind; (ii) closer link of mean daily amplitudes of both interplanetary and ground oscillations with ‘tomorrow’ values of the solar wind velocity than with current values; and (iii) correlation of the intensity of ULF waves in the solar wind, normalized to the IMF magnitude, with fluxes of relativistic electrons in the magnetosphere.