Space luminous environment adaptability of missile-borne star sensor

To solve the problem of stray interference to star point target identification while a star sensor imaging to the sky, a study on space luminous environment adaptability of missile-borne star sensor was carried out. By Plank blackbody radiation law and some astronomic knowledge, irradiancies of the stray at the star sensor working height were estimated. By relative astrophysical and mathematics knowledge, included angles between the star sensor optical axis point and the stray at any moment were calculated. The calculation correctness was verified with the star map software of Stellarium. By combining the upper analysis with the baffle suppression effect, a real-time model for space luminous environment of missile-borne star sensor was proposed. By signal-noise rate （SNR） criterion, the adaptability of missile-borne star sensor to space luminous environment was studied. As an example, a certain type of star sensor was considered when imaging to the starry sky on June 22, 2011 （the Summer Solstice） and September 20, 2011 （August 23 of the lunar year, last quarter moon） in Beijing. The space luminous environment and the adaptability to it were simulated and analyzed at the star sensor working height. In each period of time, the stray suppression of the baffle is analyzed by comparing the calculated included angle between the star sensor optical axis point and the stray with the shielded provided by system index. When the included angle is larger than the shielded angle and less than 90~, the stray is restrained by the baffle. The stray effect on star point target identification is analyzed by comparing the irradiancy of 6 magnitude star with that of the stray on star sensor sensitization surface. When the irradiancy of 6 magnitude star is 5 times more than that of the stray, there is no effect on the star point target identification. The simulation results are identicat with the actual situation. The space luminous environment of the missile-borne star sensor can be estimated real-timely by this model. The adaptability of the star sensor to space luminous environment can be analyzed conveniently. A basis for determining the relative star sensor indexes, the navigation star chosen strategy and the missile launch window can be provided.

Design of Optical System for Small Long-Life Star Sensor

In order to realize a high-precision and continuous working function of a star sensor,we propose a new optical system design.Considering the difficulty of the manufacturing process,the entire optical system uses a complicated Petzval structure.In this paper,the key design elements of the optical system applied for star sensors are presented and the most important performance parameters are given.The ground test results show that the system can maintain excellent detection performance on a near-surface atmospheric platform.This study provides an optical system design scheme for a high-precision and continuous operating star sensor,as well as the theoretical basis for future in-atmosphere and continuous star detection technology.

Application study of DSP and CPLD technology on the star sensor

On the basis of analyzing CCD signals timing and star image processing, a new design and a special parallel architeeture for improving star image processing are presented in this paper. In the design, the operation moving the data in expanded windows, ineluding the star, to the on-ehip memory of DSP is arranged in the invalid period of CCD frame signal. While the CCD saves the star image to memory, DSP processes the data in the onehip memory. This parallelism greatly improves the effieieney of processing. DSP HOLD mode and CPLD teehnology are used to make a shared memory between CCD and DSP. The five lightest stars in the star aequisition stage are aequired in only 3.5 ms. In 43μs, the data in five expanded windows ineluding stars are moved into the internal memory of DSP, and in 1.6 ms, five star eoordinates are aehieved in the star tracking stage.

Integrated Design and Accuracy Analysis of Star Sensor and Gyro on the Same Benchmark for Satellite Attitude Determination System

As an important sensor in the navigation systems,star sensors and the gyro play important roles in spacecraft attitude determination system.Complex environmental factors are the main sources of error in attitude determination.The error influence of different benchmarks and the disintegration mode between the star sensor and the gyro is analyzed in theory.The integrated design of the star sensor and the gyro on the same benchmark can effectively avoid the error influence and improves the spacecraft attitude determination accuracy.Simulation results indicate that when the stars sensor optical axis vectors overlap the reference coordinate axis of the gyro in the same benchmark,the attitude determination accuracy improves.

Study on the method of precision adjustment of star sensor

Star sensors are indispensable spatial measurement sensors for high-resolution earth observation and astronomical observations, and the demand for high measurement accuracy of satellite sensors continues to increase; thus, the star sensor optical machine adjustment error cannot be ignored. The commonly used installation error correction method cannot solely meet the precision analysis requirements. In this paper, the relationship between the optical machine installation and the star sensor measurement error is analyzed, and several common adjustment error correction methods are compared. An adjustment method for optical machines is proposed to meet the requirements of very high precision star sensors. The assembly precision requirements of the investigated very high precision star sensor are analyzed considering the whole machine, and then the optical components are controlled through optical precision adjustments to satisfy the precision requirements. Finally, through the complete machine calibration, the star sensor precision adjustment for an optical machine structure is verified. This method meets the requirements of very high precision sensors and is suitable for the precision adjustment of optical machine structures, which is of practical significance to improve the precision of star sensors.

An Approach to Star Map Simulation for Star Sensor Considering the Effect of Image Motion

To test high resolution and dynamic performance of star sensor, a method of consideration image motion on Modeling the motion blur of star sensor is proposed. Firstly, image motion geometric model based on the rotation of Starlight vector is studied. Secondly, with the help of the normal distribution of static star image energy model, introducing the star image motion speed, obtaining the energy distribution function of moving stars, implementing high dynamic simulation of star map. Finally, establishing the simulation environment, through adjusting input parameters such as integral time, rate of change of three attitude angle, the launch time, location, then, important simulation data of stars observed by star sensor in orbit can quickly be obtained, such as navigation stars information, value and direction of image motion, intensity distribution, signal to noise ratio. This work is very important to research and evaluate the star image motion compensation algorithm.

Image Smearing Modeling and Verification for Strapdown Star Sensor

To further extend study on celestial attitude determination with strapdown star sensor from static into dynamic field, one prerequisite is to generate precise dynamic simulating star maps. First a neat analytical solution of the smearing trajectory caused by spacecraft attitude maneuver is deduced successfully, whose parameters cover the geometric size of optics, three-axis angular velocities and CCD integral time. Then for the first time the mathematical law and method are discovered about how to synthesize the two formulae of smearing trajectory and the static Gaussian distribution function （GDF） model, the key of which is a line integral with regard to the static GDF attenuated by a factor 1/Ls （Ls is the arc length of the smearing trajectory） along the smearing trajectory. The dynamic smearing model is then obtained, also in an analytical form. After that, three sets of typical simulating maps and data are simulated from this dynamic model manifesting the expected smearing effects, also compatible with the linear model as its special case of no boresight rotation. Finally, model validity tests on a rate turntable are carried out, which results in a mean correlation coefficient 0.920 0 between the camera images and the corresponding model simulated ones with the same parameters. The sufficient similarity verifies the validity of the dynamic smearing model. This model, after pa- rameter calibration, can serve as a front-end loop of the ground semi-physical simulation system for celestial attitude determination with strapdown star sensor.

Accelerometer Bias and Star Sensor Installation Error Joint Calibration for the SINS/CNS Integrated Navigation System Using Attitude Maneuver

The integrated strap-down inertial nav igation system/olelestial navigation system(SINS/CNS)i an important autonomous navigation method with efective concealment and high predision.Both accelerometer biss and star ensor installation error ame important factors that aflect the performanoe of this mavigation system,which needl to be calibratexd and compensatedl.A new acelerometer bias and star sensor installation error joint calibration method for the SINS/CNS integrated navigation system i propoeed.In this newly propoeed method,the installation error of star sensor is augmented to the state vector,and the star vector,nadir angle,horkzontal poeition error and velbcity error ame ueed a8 measurementa to calbrate the two errors mentioned above.Simulations show that both accelerometer bias and star sensor installation enror an be calibratedl efectively.

Autonomous satellite constellation orbit determination using the star sensor and inter-satellite links data

A method of autonomous orbit determination for a satellite constellation using a star sensor combined with inter satellite links(ISLs) is studied.Two types of simulated observation data,Three-Satellite Constellation ISLs and background stellar observations by a CCD star sensor,are first produced.Based on these data,an observation equation is built for the constellation joint autonomous orbit determination,in which the simulations are run.The accuracy of this method with different orbital determination models are analyzed and compared with regard to the effect of potential measurement errors.The results show that autonomous satellite constellation orbit determination using star sensor measurement and ISLs data is feasible.Finally,this paper arrives at several conclusions which contribute to extending this method to a more general satellite constellation.

A method of autonomous orbit determination for satellite using star sensor

In this paper a method of autonomous orbit determination using star sensor is studied. By building relatively consummate dynamical models which simulate attitude motion of satellite and observation from satellite to background stars, the simulant computation of this method is executed, and it is shown that the method of autonomous orbit determination is feasible. Academic and calculation analyses have been done for the relation between the direction of star sensor with respect to satellite-body coordinate system and the accuracy of autonomous orbit determination.

星敏感器CMOS电路板靶面自动装调系统

CMOS电路板靶面是星敏感器结构中的关键部分,其安装姿态影响离焦距离的准确性,进而影响星敏感器成像性能。因此,靶面的装调是确保星敏感器成像质量的重要环节。当前,靶面装调主要依靠人工装调完成,存在装调精度低、产品一致性差、周期长等问题,导致产品的良品率较低。为此,研制了一台星敏感器CMOS电路板靶面自动装调设备。采取非接触式测量方式,集成测量模组与微动平台完成CMOS靶面与基准面的相对位姿测量,解决由星敏感器特殊结构造成的狭小空间内高精度测量难题。设计调整机构实现零件任意角度翻转,消除测量方向与装配方向不一致造成系统结构布置复杂的影响。最后,对测量系统进行精度分析,采用局部枚举法开发了调整垫片研磨量算法,解决由平面姿态单一已知量反求多垫片研磨量的欠定问题。实验结果表明:该系统可实现靶面自动装调功能,测量系统的重复性为1.6′,满足技术指标要求。

临近空间星敏感器观星能力研究

临近空间飞行器对星敏感器需求日益增加,与空间应用环境相比,临近空间包含大气消光、天空背景光、高速、高温和高动态等影响星敏感器观星能力的新环境.对影响临近空间星敏感器观星能力的因素进行逐级分解,得出视场、灵敏度、星点信噪比和图像处理算法等因素对观星能力的影响机理,并针对信噪比与环境因素和设计因素的关系进行量化分析,在此基础上提出临近空间星敏感器整机参数设计方法.针对特定需求开展原理样机的设计与试验验证,试验数据验证了原理样机能够满足应用需求.

一种基于虚拟星敏感器的卫星姿态确定方法

对于配置多个星敏感器的卫星,采用常规的定常增益卡尔曼滤波方法进行姿态确定时,存在滤波定常增益矩阵众多、系统复杂的问题,为了简化滤波系统设计,统一定常增益矩阵,提出一种基于虚拟星敏感器的姿态确定方法.给出一种计算量小、适合星载计算机在轨实时计算的星敏感器时间滞后补偿及相对基准标定算法,将星敏感器的输出数据统一到当前星时,同时将星敏感器的测量基准统一.基于单星敏/双星敏的输出数据构造虚拟星敏感器(安装矩阵为单位阵)的输出数据,设计统一的定常增益矩阵进行姿态确定.仿真结果表明,本方法与常规的定常增益卡尔曼滤波方法姿态确定精度相当,从而验证该方法有效,且具有重要的工程应用价值.

基于组合热流模拟策略的遥感卫星热平衡试验设计与验证

针对某型遥感卫星的不同区域对高精度热流模拟的需求,文章提出一种组合热流模拟策略:以红外加热笼来实现卫星舱板表面的外热流模拟,用薄膜电加热器实现卫星外露复杂表面外热流模拟。试验结果表明,在文述参试卫星状态及所采取的热设计措施下,该方法可对卫星低温初期、高温末期和高温热真空工况下的外热流进行精准模拟,并可有效考核遥感卫星的热控设计。该热流模拟策略可为卫星瞬态功耗散热设计的优化提供参考依据,为我国未来光学遥感卫星的热平衡试验及验证提供技术支撑。

近地轨道高精度一体式星敏感器热设计及仿真验证

某型高精度一体式星敏感器指向精度高,对温度变化非常敏感。其所处近地轨道外热流复杂多变,一体式结构和内热源集中的综合因素不仅导致散热设计困难,而且镜头直接受到内热源发热影响难以保障指向精度。首先,结合轨道参数,安装布局获得星敏感器平均吸收外热流。然后,通过分析外热流与内热源工作情况,采用被动热控和主动热控相结合的热设计方法,并对星敏感器散热面的位置与大小进行设计与计算。最后,根据轨道环境和热控措施并利用热仿真软件进行热分析验证。仿真结果表明,安装法兰温度为19.82℃~20.10℃,镜头轴向温差小于2.23℃,周向温差小于0.48℃,电路盒温度为19.10℃~23.49℃,满足热控指标。通过合理的热控设计保证了极高精度星敏感器的稳定工作条件,星敏感器的热设计合理有效。

折/衍混合光学系统在纳型星敏感器中的应用

微纳卫星是一种新兴的微小卫星,与之适配的纳型星敏感器的光学系统应在小型化的同时实现大相对孔径、高成像质量。针对这一需求,由星敏感器的系统设计精度和探测星等分配了光学系统设计参数及指标,设计了一种全视场17°、焦距25 mm、相对孔径1∶1.086、光学系统总长36.4 mm的七片式折衍混合结构的星敏感器光学系统。根据衍射元件的负色散特性,用折衍混合透镜代替常规材料正负透镜组合的方式减少了镜片数量,缩短了光学系统长度,并提高了光学系统的成像质量,使点扩散函数趋于正态分布,有利于质心提取精度的提高。所设计的光学系统在3×3像元内的能量集中度超过90%;最大畸变仅0.013%;全视场垂轴色差小于0.6μm。通过标量衍射理论计算衍射元件在520~780 nm内衍射效率可以达80%以上,通过杂散光分析软件ASAP对系统进行了视场内鬼像分析,系统在多级衍射时探测目标成像光斑路径下的光通量与最亮鬼像光斑路径下的光通量之比大于3.17×10^(4),验证了在极限星等为6.5时满足信噪比大于8.1的结论。最后对所设计的系统进行了公差分析,结果表明加工后该光学系统可满足微纳卫星对星敏感器的使用要求。

星敏感器热稳定性地面验证技术

受空间复杂热环境影响,星敏感器姿态测量结果存在轨道周期的光轴热漂移低频误差,因此需对星敏感器进行热设计并进行热稳定性试验验证。文章提出星敏感器热稳定性试验总体方案,并建立热稳定性试验系统;依据航天器真空热环境试验规范,提出星敏感器热稳定性试验判据与条件;针对试验系统误差提出控制方法并针对某型星敏感器开展热稳定性试验,结果表明星敏感器光轴热漂移量为0.114(″)/℃,满足设计指标要求。试验验证了该星敏感器热设计的有效性。

色差对星敏感器成像模型参数影响分析与补偿

星敏感器在投入使用之前需要对其进行标定以获取准确的成像模型参数值来保证姿态测量精度。现有标定方法采用单一谱段光源模拟恒星成像未考虑恒星光谱特性,色差导致标定的模型参数值和真实值存在差异。为提高模型参数准确度,首先建立了光谱差异下星敏感器恒星成像模型,进一步分析了色差对不同光谱型恒星成像模型参数的影响。在此基础上,提出采用三个典型谱段模拟恒星成像,利用三谱段成像位置信息来推算不同光谱型恒星在星敏感器工作谱段范围内成像位置,以实现对任意光谱型恒星成像模型参数的补偿。实验结果表明:相对于现有利用单一谱段标定参数测量的星角距,该补偿方法测量的星角距均方根误差降低了40.81%,有效减小了色差对星矢量测量的影响。

大口径热不敏星敏感器光机设计与分析

为满足大口径长焦距星敏感器的需求,对一种大口径热不敏星敏感器进行光机设计,根据指标要求,对热不敏系统进行光机热集成分析。使用MSC.Patran软件对主次镜结构施加温度载荷,计算出主次镜结构的热弹性变形,先利用MSC.Nastran软件计算热变形后节点的刚体位移,再利用Sigft光机接口软件分析得到变形后主次镜表面的Zernike多项式系数。将结果导入Zemax中,预判镜片面型变化以及刚体位移对弥散斑、光轴漂移量及波像差的影响。最后通过在20℃±5℃的温度范围内装调测试验证了系统性能满足指标要求以及光机热集成分析的准确性,提供了一套准确快捷的光机热集成分析流程。

星模拟器的光学系统设计研究

基于数字微镜器件(DMD)的工作原理,设计一种工作在0.4~0.7μm波段的星模拟器的光学系统,用于满足星敏感器的性能测试与评估需求。光学系统包括照明系统、棱镜、DMD和投影物镜。照明方式采用科勒照明保证光线均匀照射,其均匀性达到89%。投影系统焦距为87 mm,半视场角为5.75°,出瞳直径为31 mm,出瞳距离为33 mm。该系统经过优化,在37 lp/mm的空间频率下,MTF大于0.7,接近衍射极限,系统畸变小于1%,显示出良好的成像质量。同时,在保证投影光学系统的偏心、倾斜、厚度等公差参数较宽松的情况下,成品率高。总体上,该系统结构紧凑,性能优越,为星空模拟和测试领域带来了重要的参考。