Sun Sensor Research Articles

모멘텀 바이어스 인공위성의 2축 자세제어 시스템 설계

위성기술의 비약적 발달에 따라서 설계 및 제작에 소용되는 비용은 저렴하지만 신뢰도가 높은 인공위성 자세제어 시스템 개발이 요구되고 있다. 본 연구는 이러한 요구를 만족시키기 위해 반작용휠에 의한 모멘텀 바이어스 벡터가 임의의 방향(태양 방향)을 지향하고 안정화되는 위성시스템을 제시하였다. 위성 시스템에서 고장 가능성이 가장 적은 자기장 센서, 저정밀 태양센서 및 자기토커를 센서와 구동기로 사용하였으며, 고전적 선형 제어방법에 의해 2축 제어하는 제어시스템 설계방법을 제시하였다. 제어기는 PD 형태의 간단한 제어기가 사용되었고, 선형화된 위성시스템에 대한 PD 제어기 설계방법이 적절한 가정과 함께 제시되었다. 제시된 제어기 설계방법에 의해 설계된 폐루프 시스템의 장기 안정성 검증을 위해서 비선형 시뮬레이션 방법을 사용하였다. It is required to develop a highly reliable attitude & orbit control system of satellite that is less expensive as the technology of satellite design & integration is recently matured dramatically. To accomodate this kind of needs, the two axis attitude control method for wheel-based momentum-biased satellite system whose momentum bias vector points to a certain direction(sun direction), is developed using simple but reliable sensors and actuator: three axis magnetometer and coarse sun sensor are used as sensors, and magnetic torque bars are used as actuator. Classical PD type controller design methodologies are applied on a satellite system for the two axis control with the proper assumptions. Nonlinear simulation results are included to demonstrate the long term stability and the performance of closed-loop system design results.

Estimating near-surface air temperature with NOAA AVHRR

A technique is presented for producing estimates of near-surface air temperature (Ta) in complex terrain based on biweekly composite data from the National Oceanic and Atmospheric Administration (NOAA) advanced very high resolution radiometer (AVHRR). Results are tested against independently derived DAYMET gridded meteorological data. The model utilizes radiant surface temperature (Ts) extrapolations as a surrogate for Ta, and digital terrain information is used to adjust temperatures as a function of environmental adiabatic lapse rates. The Earth–Sun–sensor geometry information is used to constrain air temperature estimates in three different implementations. Increasing the complexity of geometric constraints on the model resulted in increased accuracies of the predictions, and yet caused a decrease in the number of days Ta estimates could be made. When comparing DAYMET surfaces to the satellite estimates of Ta, correlation coefficients (R) as high as 0.742 and standard errors of the estimate of 2.73 were observed in the final model implementation. The logic presented here serves as a useful technique for relatively simple derivation of near-surface air temperatures in a variety of remote sensing applications.

Development of a MOEMS sun sensor for space applications

This paper presents the development of a miniaturized, thin film sun sensor for sun angle detection in space applications. The overall design, different thin film layer sequences, processes, materials used and system integration are described. The sun sensor has a field of view (FoV) of greater than 2 π sr (±90°) and shall have a resolution of approximately 1° in elevation and azimuth angle. This sun sensor is therefore a coarse sun sensor with the advantage of having a large field of view. Key elements are the curved shape, the photosensitive layer consisting of copper indium gallium diselenide (CIGS), the transparent conductive layer consisting of thin molybdenum or aluminum-doped zinc oxide, and its integrated design. Different possible layer sequences are presented and discussed. The sun sensor will be one of the sensors in the attitude determination and control system (ADCS) of a nano-satellite.

A Novel Silicon Nanotips Antireflection Surface for the Micro Sun Sensor

We have developed a new technique to fabricate an antireflection surface using silicon nanotips for use on a micro Sun sensor for Mars rovers. We have achieved randomly distributed nanotips of radii spanning from 20 to 100 nm and aspect ratio of approximately 200 using a two-step dry etching process. The 30 degrees specular reflectance at the target wavelength of 1 microm is only about 0.09%, nearly 3 orders of magnitude lower than that of bare silicon, and the hemispherical reflectance is approximately 8%. When the density and aspect ratio of these nanotips are changed, a change in reflectance is demonstrated. When surfaces are covered with these nanotips, the critical problem of ghost images that are caused by multiple internal reflections in a micro Sun sensor was solved.

과학기술위성 2호(STSAT-2)의 고 정밀 디지털 태양센서(FDSS) 설계 및 분석

We have developed the FDSS (Fine Digital Sun Sensor) for the space technology of the STSAT-2 (Seience & Technology Satellite 2). The FDSS is firstly developed by using CMOS image sensor(CIS) in South Korea. The FDSS consists of the optics part, FPGA(Field Programable Gate Array) part, and MCU(Micro controller unit)part. This paper will focus on the optical characteristics of the optics part and describe the configuration of FDSS with the design of aperture. We also analyze the characteristic of optics about the pixel of the CMOS image sensor.會ĀȀ肔?⨀춙Ā搀會Ā搀𵎖⨀룎颡ĀĀĀ會ĀĀゕ?⨀耆얲ĀĀĀ會ĀĀ袕?⨀냺ꊣЀĀЀ會ĀЀ?⨀僻ꊣЀĀ؀會Ā؀㢖?⨀삙ࠀĀࠀ會Āࠀ邖?⨀룒Ā᐀會Ā᐀?⨀ᘀĀ저會Ā저䂗?⨀墚Ā᐀會Ā᐀颗?⨀Ā᐀會Ā᐀?⨀좺⒥瀀ꀏ會Ā䢘?⨀룒Ā᐀會Ā᐀ꂘ?⨀냌断

Micro-Mars: A low-cost mission to planet Mars with scientific orbiter and lander applications

The proposed Micro-Mars mission can contribute substantially to the international Mars exploration programme within the framework of a future low-cost mission. The concept consists of an orbiter integrating a total scientific payload of 30 kg including a light-weight lander of 15 kg. The spacecraft will be launched as piggyback payload by an Ariane 5 ASAP with a total launch mass of 360 kg. It will use a bipropellant propulsion system with 210 kg of fuel and four thrusters of 22 N, and four of 10 N for orbit and attitude control. Further attitude actuation shall be performed by three reaction wheels and a gyropackage, a star sensor and a sun sensor for attitude sensing. Communication will be performed in S- and X-Band to Earth and in UHF between orbiter and lander. From a highly elliptical orbit with a periapsis below 200 km, four instruments will perform high-resolution remote sensing observations and the payload consists of a camera system, a magnetometer, a dosimeter, and an ultrastable oscillator for radio science. The light-weight micro lander is a challenging technological experiment by itself. It is equipped with a suite of scientific instruments which will supplement the orbiter measurements and concentrate on the environment (temperature cycle, atmosphere, magnetosphere, and radiation).

Equal-Chord Attitude Determination Method for Spinning Spacecraft

The equal-chord method offers a straightforward and low-cost technique for the determination of the spin-axis attitude using sun and Earth sensor data. The Earth aspect angle follows from the time at which the chord lengths measured by the Earth sensor's two pencil beams are equal. An estimation technique is not required, but linear or quadratic fitting of the sensor data should be performed to remove the random errors. The accuracy of the attitude solution obtained by the equal-chord method in the presence of the relevant biases is evaluated. The result is insensitive to uniform biases in the measured chord angles, caused for instance by errors in the Earth's infrared horizon. Finally, the application of the method is demonstrated using actual flight data of the CONTOUR spacecraft.

Attitude Control System of the Wilkinson Microwave Anisotropy Probe

The Wilkinson Microwave Anisotropy Probe mission produces a map of the cosmic microwave background radiation over the entire celestial sphere by executing a fast spin and a slow precession of its spin axis about the sun line to obtain a highly interconnected set of measurements. The attitude control system implements this spin-scan observing strategy while minimizing thermal and magnetic fluctuations, especially those synchronous with the spin period. The spacecraft attitude is sensed and controlled using an inertial reference unit, 2 star trackers, a dual-head digital sun sensor, 12 coarse sun sensors, 3 reaction wheel assemblies, and a propulsion system. Sufficient attitude knowledge is provided to yield instrument pointing to a standard deviation (1σ) of 1.3 arc-min per axis. The attitude control system also maintains the spacecraft attitude during orbit maneuvers, controls the spacecraft angular momentum, and provides for safety in the event of an anomaly. An overview of the design of the attitude control system to carry out this mission is presented, as well as some early flight experience.

DEVELOPMENT OF ULTRA-LIGHT 2-AXES SUN SENSOR FOR SMALL SATELLITE

This paper addresses development of the ultra-light analog sun sensors for small satellite applications. The sun sensor is suitable for attitude determination for small satellite because of its small, light, low-cost, and low power consumption characteristics. The sun sensor is designed, manufactured and characteristic-tested with the target requirements of <TEX>${\pm}60^{\circ}$</TEX> FOV (Field of View) and pointing accuracy of <TEX>${\pm}2^{\circ}$</TEX>. Since the sun sensor has nonlinear characteristics between output measurement voltage and incident angle of sunlight, a higher order calibration equation is required for error correction. The error was calculated by using a polynomial calibration equation that was computed by the least square method obtained from the measured voltages vs. angles characteristics. Finally, the accuracies of 1-axis and 2-axes sun sensors, which consist of 2 detectors, are compared.

운용모드에 따른 과학기술위성2호의 전력 수요예측 분석

과학기술위성2호는 2007년 12월에 발사될 예정이다. 이 위성의 주관측기는 DREAM으로서, 주요 임무는 지구 또는 대기로부터 발생되는 복사에너지를 라디오파 대역에서 관측하는 것이다. 이 외 과학기술위성2호는 과학기술 실험용기기로서 정밀디지탈 태양센서, 2중헤드 별센서 등이 탑재되며, 과학기술위성2호의 자세제어 및 모멘텀 덤핑용 과학기술실험용 탑재체로서 펄스형 플라즈마 추력기가 실린다. 본 논문에서는 관측기기 및 과학기술실험용 탑재체 등의 운용모드를 고려한 과학기술위성2호의 운용모드에 따른 전력의 수요예측에 대하여 연구하였고, 임무를 수행하는 동안 안정적인 전력을 공급하기위한 필요전력에 대하여 분석하였다. STSAT-2 will be launched on December 2007 by the first Korean launch vehicle KSLV-1, and its one of the main instruments is DREAM (Dual Channel Radio Frequency and Environment Atmosphere Monitoring) which detects a signal for atmosphere from the Earth by using micro-wave signal. The STSAT-2 has many units for technology demonstration such as FDSS (Fine Digital Sun Sensor) and DHST (Dual Head Star Tracker) including PPT (Pulsed Plasma Thruster) for attitude control and momentum dumping in the space. In this paper, the power budget analysis for STSAT-2 will be studied and provided for supporting the whole mission life time during the mission of its spacecraft.

Attitude estimation from magnetometer and earth-albedo-corrected coarse sun sensor measurements

For full 3-axes attitude determination the magnetic field vector and the Sun vector can be used. A Coarse Sun Sensor consisting of six solar cells placed on each of the six outer surfaces of the satellite is used for Sun vector determination. This robust and low cost setup is sensitive to surrounding light sources as it sees the whole sky. To compensate for the largest error source, the Earth, an albedo model is developed. The total albedo light vector has contributions from the Earth surface which is illuminated by the Sun and visible from the satellite. Depending on the reflectivity of the Earth surface, the satellite's position and the Sun's position the albedo light changes. This cannot be calculated analytically and hence a numerical model is developed. For on-board computer use the Earth albedo model consisting of data tables is transferred into polynomial functions in order to save memory space. For an absolute worst case the attitude determination error can be held below 2 ∘ . In a nominal case it is better than 1 ∘ .

Aps-based miniature sun sensor for earth observation nanosatellites

This paper describes the research activity at the university of Naples concerning the design and prototype development of a low-cost APS-based miniature sun sensor for micro/nanosatellites applications. In the paper the sensor architecture is described, as well as the development of its prototype and of the facility for laboratory testing. Part of the project is also the development of a numerical code to simulate the sensor operation and predict the sensor performance. Numerical results show that an angular accuracy better than 1 arcmin is achievable.

Micro Sun Sensor for Spacecraft Attitude Control

A micro sun sensor is being developed for use on a Mars rover for the Mars ScienceLaboratory Mission. The micro sun sensor, which is basically a small pinhole camera, consists of a small mask with pinholes, placed on top of an image detector. Images of the sun are formed on the image detector when the sunilluminates the mask. Image processing is performed in the sun sensor that outputs 4 sun centroids, which the rover's main computer converts into sun angles for high gain antenna pointing and heading determination.

Precision pointing of imaging spacecraft using gyro-based attitude reference with horizon sensor updates

Remote sensing satellites are required to meet stringent pointing and drift rate requirements for imaging operations. For achieving these pointing and stability requirements, continuous and accurate three-axis attitude information is required. Inertial sensors like gyros provide continuous attitude information with better short-term stability and less random errors. However, gyro measurements are affected by drifts. Hence over time, attitudes based on the gyro reference slowly diverge from the true attitudes. On the other hand, line-of-sight (LOS) sensors like horizon sensors provide attitude information with long-term stability. Their measurements however are affected by the presence of random instrumental errors and other systematic errors. The limitations of inertial and line-of-sight sensors are mutually exclusive. Hence, by optimal fusion of attitude information from both these sensors, it is possible to retain the advantages and overcome the limitations of both, thereby providing the precise attitude information required for control.This paper describes an improved earth-pointing scheme by fusion of the three-axis attitude information from gyros and horizon sensor roll and pitch measurements along with yaw updates from the digital sun sensor. A Kalman Filter is used to estimate the three-axis attitude by online estimation and corrections of various errors from the sensor measurements. Variations in orbit rate components are also accounted for using spacecraft position and velocity measurements from the satellite positioning system. Thus precise earth-pointing is achieved

Observability of a Geosynchronous Spacecraft Attitude Determination System

The observability is examined of a standard geosynchronous spacecraft attitude determination system that uses Earth and sun sensor measurements. The results show that the attitude determination error state, including the attitude and gyro bias errors, is entirely observable over the portion of the orbit where both sun and Earth sensor measurements are available. However, when the sun is outside the sun sensor window, the system is unobservable with a one-dimensional unobservable subspace. It is shown that the eigenvalue associated with this unobservable subspaceiszero and thatthecorresponding eigenvectorisanequilibrium lineoftheattitudedetermination system. The state vector converges to this equilibrium line from any initial state. This result has practical importance in spacecraftattitudecontrolandmaybeusedtospeedtherecoveryfromattitudedeterminationanomalies.Numerical simulations and actual spacecraft e ight data are included to illustrate the theory.

A method of astronomical autonomous orbit and attitude determinations for satellites

A method of realtime autonomous orbit determination for earth satellites using the extended Kalman filtering is proposed. The observed quantities are: the satellite-sun direction vector measured by a sun sensor, the satellite-earth and satellite-moon direction vectors measured by an ultraviolet sensor, and the geocentric distance measured by a radar altimeter. At the same time the satellite attitude to the earth is also determined. Results of our simulation of the autonomous orbit determination show that the precision of the orbit determinations is better than 200 m. The effects of the sampling period, orbital inclination, orbital eccentricity and orbital altitude on the precision of orbit determination are analyzed and compared, and certain principles helpful for improving the precision of orbit determination are suggested.

BILSAT-1: A low-cost, agile, earth observation microsatellite for Turkey

TUBITAK-BIlTEN has initaited a project to develop and propagate small satallite technologies in Turkey.As part of this initiative, TUBITAK-BILTEN is working working SSTL to develop a 100kg class enhanced microsatelliote, BILSAT-1. With the successful completion of this project, TUBITAK-BILTEN will be capable of producing its own satellites, covering all phases from design from design to production and in-orbit operatiion. It is hoped that acquisition of these technologies will stimulate Turkish industry into greater involvement in space related activities. The project was started in August 2001 and will run through to February 2003 with launch scheduled for July 2003. BILSAT-1 will be one of the most capable satellites that SSTL have eveer built and features several technologies normally only found on larger satellites. Specifically, the Attitude Determination and Control System of BILSAT-1 will be the most advanced that SSTL have everflown: Dual redundant Star Cameras, sun sensors and rate gyros provide accurate and precise attitude information allowing a very high degree of attitude knowledge. Acruators on board will make the satellite extremely agile — for instance allowing fast slew manoeuvers about its roll and pitch axes. The agile control system also enables ground target revisit times to be reduced compared to nadir-pointing gravity gradient stabilized satellites, and will allow stereoscopic imaging, target tracking and multiple attitude imaging to be undertaken with the satellites prime payloads: a 4-band multispectral 26-metre GSD imaging system and a 12-metre GSD panchromatic imager. Also on board the satellite there are additional payloads, including a state-of-the-art Digital Signal Processing payload (GEZGIN) that will enabele real time image compression in JPEG2000 format using a high performance floating point DSP, and a low resolution multi spectral (9-band) camera (COBAN). BILSAT-1 will also co-operate in the international Disaster Monitoring Constellation (DMC) led by SSTL, providing the ability to enhance the imaging capabilities of the constellation. In parallel with the microsatellite design and build activities, all the infrastructure required to design, produce and operate a satellite is being constructed at BILTEN's premises in Turkey. This infrastructure includes assembly and integration rooms, a PCB prototyping workshop, research and development laboratories, and a satellite mission control ground station.

Champ–The first flexbus in orbit

The launch of the German research satellite Champ was the successful premiere for a new ASTRIUM GmbH satellite concept. The Flexbus satellite design and management concept was developed in view of shrinking budgets and great demand on reliable satellites enabling affordable missions in the area of science and earth observation. Compared to a conventional design and management approach the Flexbus concept applied on the Champ satellite development resulted in a cost reduction by factor 3. The use of a mature basic architecture for the spacecraft bus and the combination of proven components with mission-specific payload elements make Champ a satellite which optimally adapts to individual mission and payload requirements. Also the cooperation with the customer and principal investigator GeoForschungsZentrum Postdam (GFZ) and the satellite operator DLR was based on a new approach. Both partners were involved very early in the design and development phase. Besides the proven technologies Champ features a first in orbit use of the in-house developed Coarse Earth and Sun Sensor CESS, foam isolated body fixed solar arrays and a low cost boom deployment mechanism. All three new developments performed very well.

3315 アラスカ大学との共同学生ロケット 2 号機の打ち上げ実験

Our Tokai student rocket project team (TSRP) has launched a collaborative rocket with University of Alaska Fairbanks (UAF) and Toyama Prefectural University (TPU) at Alaska on 18,March 2002. Except the motor section, all the parts of this rocket were completely made by students. After we designed and constructed a tri-axial fluxgate magnetometer and a sun sensor, we have tested various environmental experiments and assembled the payload section at UAF. The collaborative rocket that was 5.4 meter of length has been reached 89km of altitude and we got good data from our instruments. We show here our instruments and an experimental result of the sounding rocket observaation.

Complex Modelling – An Important Means for Performance Prediction of Spacecraft Equipment

A mathematical model describes the behaviour of a real system with a defined accuracy and a defined scope using mathematical methods. A complex model is a model that contains the combination of different functional principles, like optics and electronics. An advanced language of technical computing integrates computation, visualisation and programming and allows an effective interactive operation for developing and using models. The development of complex models for studying the product behaviour and for predicting special effects during design and operation is presented by means of the examples of a four-quadrant sun sensor, a laser beam deflection system and the camera function of a rendezvous and docking videometer.