Effects of in-track maneuver on Sun illumination conditions of near-circular low Earth orbits

By developing approximate analytical models considering the J₂ perturbation, the effects of an in-track maneuver on the orbital Sun illumination conditions of near-circular low Earth orbits are analyzed. First, two approximate models for the variations in orbital sunshine angles are developed, one for variations at a given time and the other for variations at a given argument of latitude. Next, two approximate models for variations in orbital arc in Earth shadow are developed, one considers the small eccentricity and the other uses the zero eccentricity. Finally, the developed approximate models are applied to analyzing the Sun illumination conditions of a typical in-track maneuver mission on a near-circular low Earth orbit. From the results obtained, three major conclusions can be drawn. First, the variations in orbital sunshine angles at a given time may reach tens of degrees when the drifting time reaches hundreds of orbital periods, and the approximate model for that situation cannot effectively approach the numerical results. Second, the variations in orbital sunshine angles for any given argument of latitude are only a couple of degrees even when the drifting time reaches 500 orbital periods, and the approximation model developed can effectively approach the numerical results. Third, for variations in orbital arc in Earth shadow, the approximate model considering the small eccentricity has simple expressions and can effectively approach the numerical results; in contrast, the approximate model using the zero eccentricity has relatively worse precision.     

Investigation of dynamics of discrete framed structures by a numerical wave-based method and an analytical homogenization approach

In this article, the analytical homogenization method of periodic discrete media (HPDM) and the numerical condensed wave finite element method (CWFEM) are employed to study the lon-gitudinal and transverse vibrations of framed structures. The valid frequency range of the HPDM is re-evaluated using the wave propagation feature identified by the CWFEM. The relative error of the wavenumber by the HPDM compared to that by the CWFEM is illustrated in functions of fre-quency and scale ratio. A parametric study on the thickness of the structure is carried out where the dispersion relation and the relative error are given for three different thicknesses. The dynamics of a finite structure such as natural frequency and forced response are also investigated using the HPDM and the CWFEM.     

Determination of station positions and velocities from laser ranging observations to Ajisai,Starlette and Stella satellites

The positions and velocities of the four Satellite Laser Ranging (SLR) stations: Yarragadee (7090), Greenbelt (7105), Graz (7839) and Herstmonceux (7840) from 5-year (2001–2005) SLR data of low orbiting satellites (LEO): Ajisai, Starlette and Stella were determined. The orbits of these satellites were computed from the data provided by 20 SLR stations. All orbital computations were performed by means of NASA Goddard’s GEODYN-II program. The geocentric coordinates were transformed to the topocentric North–South, East–West and Vertical components in reference to ITRF2005. The influence of the number of normal points per orbital arc and the empirical acceleration coefficients on the quality of station coordinates was studied. To get standard deviation of the coordinates determination lower than 1 cm, the number of the normal points per site had to be greater than 50. The computed positions and velocities were compared to those derived from LAGEOS-1/LAGEOS-2 data. Three parameters were used for this comparison: station coordinates stability, differences from ITRF2005 positions and velocities. The stability of coordinates of LEO satellites is significantly worse (17.8 mm) than those of LAGEOS (7.6 mm), the better results are for Ajisai (15.4 mm) than for Starlette/Stella (20.4 mm). The difference in positions between the computed values and ITRF2005 were little bit worse for Starlette/Stella (6.6 mm) than for LAGEOS (4.6 mm), the results for Ajisai were five times worse (29.7 mm) probably due to center of mass correction of this satellite. The station velocities with some exceptions were on the same level (≈1 mm/year) for all satellites. The results presented in this work show that results from Starlette/Stella are better than those from Ajisai for station coordinates determination. We can applied the data from LEO satellites, especially Starlette and Stella for determination of the SLR station coordinates but with two times lower accuracy than when using LAGEOS data.     

氢对TC4钛合金电子束焊接头疲劳裂纹扩展速率的影响

采用充氢CT试样对TC4钛合金电子束焊接头的室温疲劳裂纹扩展速率进行了测定,并对试样断口和接头各区显微组织进行了观察.试验结果表明:充氢母材试样在低速扩展区和失稳扩展区的da/dN相对于未充氢试样有明显的提高,但不同氢含量之间差别不大,在稳态扩展区(Paris区),氢对da/dN的影响很小.充氢焊缝试样在整个裂纹扩展过...     

CubeSail: A low cost CubeSat based solar sail demonstration mission

CubeSail is a nano-solar sail mission based on the 3U CubeSat standard, which is currently being designed and built at the Surrey Space Centre, University of Surrey. CubeSail will have a total mass of around 3 kg and will deploy a 5 × 5 m sail in low Earth orbit. The primary aim of the mission is to demonstrate the concept of solar sailing and end-of-life de-orbiting using the sail membrane as a drag-sail. The spacecraft will have a compact 3-axis stabilised attitude control system, which uses three magnetic torquers aligned with the spacecraft principle axis as well as a novel two-dimensional translation stage separating the spacecraft bus from the sail. CubeSail’s deployment mechanism consists of four novel booms and four-quadrant sail membranes. The proposed booms are made from tape-spring blades and will deploy the sail membrane from a 2U CubeSat standard structure. This paper presents a systems level overview of the CubeSat mission, focusing on the mission orbit and de-orbiting, in addition to the deployment, attitude control and the satellite bus.     

Feasibility of performing space surveillance tasks with a proposed space-based optical architecture

Under ESA contract an industrial consortium including Aboa Space Research Oy (ASRO), the Astronomical Institute of the University of Bern (AIUB), and the Dutch National Aerospace Laboratory (NLR), proposed the observation concept, developed a suitable sensor architecture, and assessed the performance of a space-based optical (SBO) telescope in 2005. The goal of the SBO study was to analyse how the existing knowledge gap in the space debris population in the millimetre and centimetre regime may be closed by means of a passive optical instrument. The SBO instrument was requested to provide statistical information on the space debris population in terms of number of objects and size distribution. The SBO instrument was considered to be a cost-efficient with 20 cm aperture and 6° field-of-view and having flexible integration requirements. It should be possible to integrate the SBO instrument easily as a secondary payload on satellites launched into low-Earth orbits (LEO), or into geostationary orbit (GEO). Thus the selected mission concept only allowed for fix-mounted telescopes, and the pointing direction could be requested freely. Since 2007 ESA focuses space surveillance and tracking activities in the Space Situational Awareness (SSA) preparatory program. Ground-based radars and optical telescopes are studied for the build-up and maintenance of a catalogue of objects. In this paper we analyse how the proposed SBO architecture could contribute to the space surveillance tasks survey and tracking. We assume that the SBO instrumentation is placed into a circular sun-synchronous orbit at 800 km altitude. We discuss the observation conditions of objects at higher altitude, and select an orbit close to the terminator plane. A pointing of the sensor orthogonal to the orbital plane with optimal elevation slightly in positive direction (0° and +5°) is found optimal for accessing the entire GEO regime within one day, implying a very good coverage of controlled objects in GEO, too. Simulations using ESA’s Program for Radar and Optical Observation Forecasting (PROOF) in the version 2005 and a GEO reference population extracted from DISCOS revealed that the proposed pointing scenario provides low phase angles together with low angular velocities of the objects crossing the field-of-view. Radiometric simulations show that the optimal exposure time is 1–2 s, and that spherical objects in GEO with a diameter of below 1 m can be detected. The GEO population can be covered under proper illumination nearly completely, but seasonal drops of the coverage are possible. Subsequent observations of objects are on average at least every 1.5 days, not exceeding 3 days at maximum. A single observation arc spans 3° to 5° on average. Using a simulation environment that connects PROOF to AIUB’s program system CelMech we verify the consistency of the initial orbit determination for five selected test objects on subsequent days as a function of realistic astrometric noise levels. The initial orbit determination is possible. We define requirements for a correlator process essential for catalogue build-up and maintenance. Each single observation should provide an astrometric accuracy of at least 1”–1.5” so that the initially determined orbits are consistent within a few hundred kilometres for the semi-major axis, 0.01 for the eccentricity, and 0.1° for the inclination.     

基于ABAQUS的渐开线齿轮齿根裂纹扩展仿真

齿轮在传动过程中经常发生断裂,严重影响了齿轮的使用。本文研究了齿轮齿根裂纹扩展特性,为齿轮安全设计打下基础。通过Pro/E参数化建模建立一对啮合渐开线齿轮,加载一定载荷后通过分析得到最大应力区域、假定裂纹源头和初始裂纹方向;在应力强度因子的计算过程中,从线弹性断裂力学角度出发,应用ABAQUS软件计算得到裂纹尖端应力强度因子;根据最大周向应力法,计算了裂纹失稳后的扩展角,并在ABAQUS中分步模拟了裂纹的扩展趋势;当kI值发生突变后,裂纹迅速扩展以致轮齿断裂。     

一种提高卫星相对轨道运动全物理仿真逼真度的新方法

现有的卫星控制系统全物理仿真很少对卫星绕地球的轨道运动进行模拟,即使在卫星间相对轨道运动的全物理仿真中也没有考虑地心引力差和惯性力项的存在,因此其逼真度受到了影响.提出一种在共面圆轨道近距离卫星相对运动全物理仿真中引入地心引力差和惯性力项的方法,提高了物理仿真实验的逼真度.     

空间实验室低轨高压长寿命半刚性太阳电池翼技术研究

介绍了国内首次使用的半刚性太阳电池翼的任务与技术指标和产品组成。给出了编织用玻璃纤维纱线制备、柔性玻璃纤维编织、半刚性基板组合工艺和低轨空间环境适应性等关键技术。阐述了研制中进行的运动学和动力学、模态、热、在轨栽荷、强度等仿真分析,以及噪声、振动、高低温展开、模态和其他试验。     

美国AEHF军事通信卫星推进系统及其在首发星上的应用

美国先进极高频（AEHF）项目将采用4颗运行于地球同步轨道（GEO）的AE—HF军事通信卫星。AEHF卫星的推进系统由远地点发动机、姿态控制发动机组和双模式霍尔电推进子系统组成。首发星AEHF-1是世界上首个采用霍尔电推进系统执行发射后轨道提升任夯的GEO卫星。AEHF-1卫星星箭分离后,远地点发动机未能正常工作,因此...