Sun Sensor Research Articles

Radiometric normalization for change detection in peatlands: a modified temporal invariant cluster approach

Radiometric normalization is a vital stage in any change detection study due to the complex interactions of radiance and irradiance between the Earth's surface and atmosphere. Compensation for variables such as sun's angle, surface profile, atmospheric conditions, and sensor calibration coefficients are essential in achieving a radiometrically stable data base of multi-temporal, multi-spectral imagery for a change detection study. In this study, five Landsat Enhanced Thematic Mapper Plus (ETM+) images taken over the east coast of Ireland in 2001 were geometrically corrected and topographically normalized for further processing and analysis. Assessment of various vegetation indices showed that the enhanced vegetation index 2 (EVI2) gave the highest accuracy in identifying the various vegetation types and habitats in the Wicklow Mountains National Park. The initial analysis of radiometric normalization with temporal invariant clusters (TICs) gave poor results due to the spectral heterogeneity of urban pixels within each image. A revised TIC subset normalized method was developed using regional growth parameters in urban environments to limit the spatial and spectral extent of pixels used in the TIC scene normalization process. Correlation analysis between the TIC-subset-normalized ETM+ data and Landsat Ecosystem Disturbance Adaptive Processing System (LEDAPS) absolute corrected data produced coefficient of determination (R²) values between 0.88 and 0.98. Such results demonstrated the robustness of the TIC subset normalization procedure when correcting for atmospheric variability between images while maintaining spectral integrity. Statistical analysis on master slave and TIC-subset-normalized slave data using cumulative distribution curves derived from image histograms showed an 86.93% reduction in the maximum difference between master and slave data due to the TIC subset normalization process. This procedure of radiometric normalization is suitable in landscapes with a low density of spectrally stable targets.

航天太阳敏感器的应用与发展

The principle and optical design of digital sun sensors widely used in different fields are reviewed and their applications and development are given.Firstly,it describes the working principle of sun sensors.The digital sun sensors are constructed with the principle of pinhole camera,and it consisted of a optical system,a photo detector and a information processing unit.Then,the paper gives the optical systems,which may be implemented with a single aperture,or a single split,or multiple apertures,or multiple splits.Furthermore,it focuses on the development of the digital sun sensors based on conventional photo detectors and micro detectors integrating the optical system,image sensor and the information processing unit,and also introduces centroiding algorithms for sun spots.Finally,the future challenge and development trends for space sun sensors are discussed.

In-flight Calibration of Attitude Determination Systems for Information Mini-satellites

Contemporary information mini-satellites have general mass up to 500 kg and are placed onto the orbit altitudes from 500 up to 1500 km. For such spacecraft some principle problems on precise attitude determination and in-flight calibration are considered and elaborated methods are presented for their solving. Original multiple discrete algorithms for filtering, integration and calibration of a strapped-down inertial navigation system for precise determining the spacecraft attitude, are presented. This system contains electrostatic micro-accelerometer, an inertial measurement unit based on the gyro sensors, an astronomical system based on star trackers, multi-head Sun sensor and magnetic sensor, the that are fixed to the spacecraft body.

A Low-Cost Photodiode Sun Sensor for CubeSat and Planetary Microrover

This paper presents the development of low-cost methodologies to determine the attitude of a small, CubeSat-class satellite and a microrover relative to the sun's direction. The use of commercial hardware and simple embedded designs has become an effective path for university programs to put experimental payloads in space for minimal cost, and the development of sensors for attitude and heading determination is often a critical part. The development of two compact and efficient but simple coarse sun sensor methodologies is presented in this research. A direct measurement of the solar angle uses a photodiode array sensor and slit mask. Another estimation of the solar angle uses current measurements from orthogonal arrays of solar cells. The two methodologies are tested and compared on ground hardware. Testing results show that coarse sun sensing is efficient even with minimal processing and complexity of design for satellite attitude determination systems and rover navigation systems.

An Autonomous Microdigital Sun Sensor by a CMOS Imager in Space Application

Digital sun sensors are a category of attitude sensors applied on spacecraft. The proposed microdigital sun sensor (μDSS) is specifically applied for microsatellites. This application requires low power consumption, high accuracy, and radiation hardness, as well as miniaturization. The μDSS adopts a low-power windowing readout method: a two-step acquisition-tracking operation. The power consumed in the acquisition mode is significantly reduced with a profiling method which is enabled by a specific pixel design. The high accuracy is achieved by reducing the system thermal noise with a quadruple sampling method. The radiation hardness is strengthened by both hardware design and algorithm correction. The system miniaturization is realized by a SoC solution: integrating the active pixel sensor, timing control circuit, algorithm and processing circuit, and analog-to-digital converter on chip. The proposed μDSS achieves a 0.01° accuracy with 21 mW at 10 frame/s.

Digital Sun Sensor Multi-Spot Operation

The operation and test of a multi-spot digital sun sensor for precise sun-line determination is described. The image forming system consists of an opaque mask with multiple pinhole apertures producing multiple, simultaneous, spot-like images of the sun on the focal plane. The sun-line precision can be improved by averaging multiple simultaneous measures. Nevertheless, the sensor operation on a wide field of view requires acquiring and processing images in which the number of sun spots and the related intensity level are largely variable. To this end, a reliable and robust image acquisition procedure based on a variable shutter time has been considered as well as a calibration function exploiting also the knowledge of the sun-spot array size. Main focus of the present paper is the experimental validation of the wide field of view operation of the sensor by using a sensor prototype and a laboratory test facility. Results demonstrate that it is possible to keep high measurement precision also for large off-boresight angles.

Research on Satellite Multi-Sensor Attitude Determination System

An attitude determination system has been designed for the geocentric pointing triaxial stabilized satellites which employ a continuously running inertial rate sensor in conjunction with sun sensor and earth sensor. Earth/sun sensor data are processed to generate corrections to satellite attitude, gyro constant drift and earth sensor drift bias estimates. An extended Kalman filter based on the attitude determination system is derived in this paper for the satellite using two earth sensors, a two-axis digit sun sensor as attitude sensors and a three-axis gyro for the angular velocity. A simulation model is developed to estimate the attitude determination performance. Simulation results show that precision attitude determination is achieved using the selected attitude hardware and algorithms.

THE STUDY ON HEIGHT INFORMATION EXTRACTION OF CULTURAL FEATURES IN REMOTE SENSING IMAGES BASED ON SHADOW AREAS

Abstract. Cultural feature is important element in geospatial information library and the height information is important information of cultural features. The existences of the height information and its precision have direct influence over topographic map, especially the quality of large-scale and medium-scale topographic map, and the level of surveying and mapping support. There are a lot of methods about height information extraction, in which the main methods are ground survey (field direct measurement) spatial sensor and photogrammetric ways. However, automatic extraction is very tough. This paper has had an emphasis on segmentation algorithm on shadow areas under multiple constraints and realized automatic extraction of height information by using shadow. Binarization image can be obtained using gray threshold estimated under the multiple constraints. On the interesting area, spot elimination and region splitting are made. After region labeling and non-shadowed regions elimination, shadow area of cultural features can be found. Then height of the cultural features can be calculated using shadow length, sun altitude angle, azimuth angle, and sensor altitude angle, azimuth angle. A great many of experiments have shown that mean square error of the height information of cultural features extraction is close to 2 meter and automatic extraction rate is close to 70%.

Error Compensation for Area Digital Sun Sensor

Compared to the error factors of the Linear Array Digital Sun Sensor (DSS), those of the Area Array DSS are complicated and methods used for error compensation are not valid or simple enough. This paper presents the main error factors of the Area Array DSS and proposes an effective method to compensate them. The procedure of error compensation of Area Array DSS includes three steps. First, the geometric error of calibration is compensated; second, the coordinate map method is used to compensate the error caused by optical refraction; third, the high order polynomial-fitting method is applied to calculate the tangent of the sun angles; finally, the arc tangent method is used to calculate the sun angles. Experimental results of the product of the High Accuracy Sun Sensor indicate that the precision is better than 0.02° during the cone field of view (CFOV) of 10°, and the precision is better than 0.14° during the CFOV 10° to 64°. The proposed compensation method effectively compensates the major error factors and significantly improves the measure precision of the Area APS DSS.

Development of a Fine Digital Sun Sensor for STSAT-2

Satellite devices for fine attitude control of the Science & Technology Satellite-2 (STSAT-2). Based on the mission requirements of STSAT-2, the conventional analog-type sun sensors were found to be inadequate, motivating the development of a compact, fast and fine digital sun sensor (FDSS). The FDSS uses a CMOS image sensor and has an accuracy of less than 0.03degrees, an update rate of 5Hz and a weight of less than 800g. A pinhole-type aperture is substituted for the optical lens to minimize its weight. The target process speed is obtained by utilizing the Field Programmable Gate Array (FPGA), which acquires images from the CMOS sensor, and stores and processes the image data. The sensor accuracy is maintained by a rigorous centroid algorithm. This paper describes the FDSS designs, realizations, tests and calibration results.

Accurate and Wide-Field-of-View MEMS-Based Sun Sensor for Industrial Applications

This paper describes the design, fabrication, simulation, and experimental results of an improved miniaturized two-axis sun sensor for industrial applications, created by adapting a technology used previously in satellite applications. The sensor for each axis is composed of six photodiodes integrated in a crystalline-silicon substrate and a layer of cover glass, which is used to protect the silicon and to hold the windows. The high precision is obtained by the subdivision of the field of view, which is ±60°, with a resolution of 0.1°. Each region is controlled by an independent high-precision solar sensor. The sensor can be used for sun-tracking applications in a photovoltaic system, heliostat concentration plants, and lighting applications.

Rigid Body Attitude Control Using a Single Vector Measurement and Gyro

Most existing methods for satellite attitude control assume that full knowledge of satellite attitude is available by algebraic manipulation of at least two vector measurements from attitude sensors such as Sun, Earth-horizon, Earth-magnetic or star tracker sensors. Kalman filtering is usually used when only one vector measurement is available, however, asymptotic stability of nonlinear and uncertain dynamics of satellite is not guaranteed in this case. This technical note presents a coupled nonlinear estimator-controller for satellite attitude determination and control by using a 3-axis gyro and a single vector measurement for a fully actuated satellite. We assume that the moment-of-inertia matrix of satellite is unknown. Asymptotic convergence of the satellite attitude and angular velocity to their desired values is proven. Performance of the proposed controller is illustrated in a realistic case study by simulation where a magnetometer is used to provide the single vector measurement.

Design and on-orbit performance of the attitude determination and control system for the ZDPS-1A pico-satellite

The ZDPS-1A pico-satellite, developed by the Zhejiang University, is featured with a three-axis stabilizing capability. It is 15×15×15cm3 cube-shaped satellite with a total mass of 3.5kg. ZDPS-1A is the first pico-satellite that has been launched successfully in China. The mission of ZDPS-1A is on-orbit system verification of student-build pico-satellite and wide range earth observation with a micro panoramic camera. A miniature momentum wheel is employed to offer gyro stiffness stability in the pitch (orbit normal) axis. Magnetic coils are employed to generate control torques to achieve the three-axis stabilization of nadir-pointing. The attitude sensors employed in the design include two three-axis magnetometers (TAMs), a three-axis gyro, and two sun sensors. Both ground simulations and on-orbit testing are conducted to verify the feasibility of the given attitude determination and control system (ADCS).

Controle em três eixos para aquisição de atitude por satélite universitário partindo de condições iniciais desfavoráveis

A concepção de um satélite universitário de baixo custo busca atender as restrições de seu lançamento como carga útil secundária. A configuração girostato foi investigada para avaliar o desempenho do Sistema de Controle de Atitude (SCA) autônomo durante a aquisição de atitude quando o satélite se separa do veículo lançador em uma condição inicial desfavorável. A configuração girostato é controlada em três eixos e busca manter voltada para o Sol a única face com o painel solar. Uma roda de momentum, cujo eixo de rotação é ortogonal àquela face, provê rigidez giroscópica para, em conjunto com três bobinas de torque com núcleo de ar, gerar torques para controle de atitude em três eixos. O SCA inicia sua ação a partir da separação, operando em malha fechada com a estimação da atitude e da velocidade angular do satélite. Essas estimativas são fornecidas por um filtro de Kalman estendido (FKE) que processa medidas vetoriais da direção do Sol e do campo geomagnético oriundas de sensores de Sol e magnetômetro triaxial. Partindo de condições desfavoráveis, os resultados mostraram que o apontamento do satélite demanda cerca de um dia e meio dia para prover erro de apontamento inferior a 10º e menos de dois dias para manter o erro abaixo de 0,1º. Em torno do eixo da roda, ocorre oscilação do satélite em torno da referência, mas dentro de uma margem aceitável. A duração das manobras em relação à capacidade de carga das baterias embarcadas confirmou a viabilidade da configuração girostato. O desempenho e a complexidade dessa configuração foram comparados com a estabilização de corpo rígido com rotação spin visando selecionar configuração apropriada para a missão do satélite.

The attitude determination system of the RAX satellite

The Radio Aurora Explorer (RAX) is a triple CubeSat that launched on November 19, 2010. RAX was designed to study plasma irregularities in the polar lower ionosphere (80–300km), and is the first CubeSat mission funded by the United States National Science Foundation. The scientific mission requires attitude knowledge within 5° (1−σ), and a custom attitude determination subsystem was developed for the mission. The subsystem utilizes rate gyros, magnetometers, coarse sun sensors, and an extended Kalman filter, and was designed to be a simple, low cost solution to meet the attitude determination requirements. In this paper, we describe the design, implementation, and testing of the RAX attitude determination subsystem, including derivation of the determination requirements, sensor selection, the integrated hardware design, pre-flight sensor calibration, and attitude estimation algorithms. The paper is meant to serve as a resource for others in the small satellite and nanosatellite communities, as well as a critical reference for those analyzing RAX data. Lessons learned from the design and performance of the RAX determination subsystem will be used in future designs of attitude determination systems for small satellites and similar platforms, such as high altitude balloons and autonomous aerial vehicles.

Field testing of visual odometry aided by a sun sensor and inclinometer

AbstractIn this paper, we present a novel approach to planetary rover localization that incorporates sun sensor and inclinometer data directly into a stereo visual odometry pipeline. Utilizing the absolute orientation information provided by the sun sensor and inclinometer significantly reduces the error growth of the visual odometry path estimate. The measurements have very low computation, power, and mass requirements, providing localization improvement at nearly negligible cost. We describe the mathematical formulation of error terms for the stereo camera, sun sensor, and inclinometer measurements, as well as the bundle adjustment framework for determining the maximum likelihood vehicle transformation. Extensive results are presented from experimental trials utilizing data collected during a 10‐km traversal of a Mars analogue site on Devon Island in the Canadian high Arctic. We also illustrate how our approach can be used to reduce the computational burden of visual odometry for planetary exploration missions. © 2012 Wiley Periodicals, Inc.

Assessing attitude error of FORMOSAT-3/COSMIC satellites and its impact on orbit determination

An attitude determination and control system (ADCS) is critical to satellite attitude maneuvers and to the coordinate transformation from the inertial frame to the spacecraft frame. This paper shows specific sensors in the ADCS of the satellite mission FORMOSAT-3/COSMIC (F3/C) and the impact of the ADCS quality on orbit accuracy. The selection of main POD antenna depends on the beta angles of the different F3/C satellites (for FM2 and FM4) during the inflight phase. In particular, under the eclipse, alternative attitude sensors are activated to replace the Sun sensors, and such a sensor change leads to anomalous GPS phase residuals and a degraded orbit accuracy. Since the nominal attitude serves as a reference for ADCS, the 3-dimensional attitude-induced errors in reduced dynamic orbits over selected days in 2010 show 9.35, 10.78, 4.97, 5.48, 7.18, and 6.89cm for FM1–FM6. Besides, the 3-dimensional velocity errors induced by the attitude effect are 0.10, 0.10, 0.07, 0.08, 0.09, and 0.10 for FM1–FM6. We analyze the quality of the observed attitude transformation matrix of F3/C and its impact on kinematic orbit determination. With 249days of GPS in 2008, the analysis leads to the following averaged 3-dimensional attitude-induced orbit errors: 2.72, 2.62, 2.37, 1.90, 1.70, and 1.99cm for satellites FM1–FM6. Critical suggestions of geodetic payloads for the follow-on mission of F3/C are presented based on the current result.

Dual‐EKF‐Based Real‐Time Celestial Navigation for Lunar Rover

A key requirement of lunar rover autonomous navigation is to acquire state information accurately in real‐time during its motion and set up a gradual parameter‐based nonlinear kinematics model for the rover. In this paper, we propose a dual‐extended‐Kalman‐filter‐ (dual‐EKF‐) based real‐time celestial navigation (RCN) method. The proposed method considers the rover position and velocity on the lunar surface as the system parameters and establishes a constant velocity (CV) model. In addition, the attitude quaternion is considered as the system state, and the quaternion differential equation is established as the state equation, which incorporates the output of angular rate gyroscope. Therefore, the measurement equation can be established with sun direction vector from the sun sensor and speed observation from the speedometer. The gyro continuous output ensures the algorithm real‐time operation. Finally, we use the dual‐EKF method to solve the system equations. Simulation results show that the proposed method can acquire the rover position and heading information in real time and greatly improve the navigation accuracy. Our method overcomes the disadvantage of the cumulative error in inertial navigation.

Magnetospheric multiscale (MMS) mission attitude ground system design

This paper describes the attitude ground system (AGS) design to be used for support of the Magnetospheric MultiScale (MMS) mission. The AGS exists as one component of the mission operations control center. It has responsibility for validating the onboard attitude and accelerometer bias estimates, calibrating the attitude sensors and the spacecraft inertia tensor, and generating a definitive attitude history for use by the science teams. NASA's Goddard Space Flight Center (GSFC) in Greenbelt, Maryland is responsible for developing the MMS spacecraft, for the overall management of the MMS mission, and for mission operations. MMS is scheduled for launch in 2014 for a planned two-year mission. The MMS mission consists of four identical spacecraft flying in a tetrahedral formation in an eccentric Earth orbit. The relatively tight formation, ranging from 10 to 400 km, will provide coordinated observations giving insight into small-scale magnetic field reconnection processes. By varying the size of the tetrahedron and the orbital semi-major axis and eccentricity, and making use of the changing solar phase, this geometry allows for the study of both bow shock and magnetotail plasma physics, including acceleration, reconnection, and turbulence. The mission divides into two phases for science; these phases will have orbit dimensions of 1.2 x 12 Earth radii in the first phase and 1.2x25 Earth radii in the second in order to study the dayside magnetopause and the nightside magnetotail, respectively. The orbital periods are roughly one day and three days for the two mission phases. Each of the four MMS spacecraft will be spin stabilized at 3 revolutions per minute (rpm), with the spin axis oriented near the ecliptic north pole but tipped approximately 2.5 deg towards the Sun line. The main body of each spacecraft will be an eight-sided platform with diameter of 3.4 m and height of 1.2 m. Several booms are attached to this central core: two axial booms of 14.9 m length, two radial magnetometer booms of 5 m length, and four radial wire booms of 60 m length. Attitude and orbit control will use a set of axial and radial thrusters. A four-head star tracker and a slit-type digital Sun sensor (DSS) provide input for attitude determination. In addition, an accelerometer will be used for closed-loop orbit maneuver control. The primary AGS product will be a daily definitive attitude history. Due to power limitations, the star tracker and accelerometer data will not be available at all times. However, tracker data from at least 10 percent of each orbit and continuous DSS data will be provided. An extended Kalman filter (EKF) will be used to estimate the three-axis attitude (i.e., spin axis orientation and spin phase) and rotation rate for all times when the tracker data is valid. For other times, the attitude is generated by assuming a constant angular momentum vector in the inertial frame. The DSS sun pulse will provide a timing signal to maintain an accurate spin phase. There will be times when the Sun is occulted and DSS data is not available. If this occurs at the start or end of a definitive attitude product, then the spin phase will be extrapolated using the mean rate determined by the EKF.

Spin-axis estimation of the Radiation Belt Storm Probes spacecraft using RF Doppler data

A methodology for using RF Doppler data for spin-axis estimation is presented. The NASA Radiation Belt Storm Probes (RBSP) mission, a dual-spacecraft mission planned for a 2012 launch, will yield valuable insights into the physical dynamics and processes of the Earth's radiation belts, particularly those that produce hazardous space weather effects and those that can affect solar system exploration. Both spacecraft have on-board Sun sensors that, in conjunction with magnetometer data, provide the required data to determine the attitude. However, the quality of the magnetometer measurements is compromised at higher altitudes because of a weaker magnetic field and also during solar storms because of variations in the Earth's magnetic field. Each spacecraft's on-board S-band radio frequency (RF) system includes a JHU/APL Frontier Radio transceiver, which provides a coherent downlink signal used for Doppler navigation. The RF system also includes two low-gain antennas offset from the spin axis with boresights parallel to the spin and anti-spin axes. Due to the spin of the spacecraft, a Doppler-induced modulation is present on the downlink carrier. Once the orbit determination is performed and the Doppler signal is analyzed, this modulation, or spin signature, is apparent. This spin signature can be used to accurately determine the spin rate, spin phase, and, in conjunction with Sun sensor data, the spin axis orientation. This provides a level of system redundancy with the magnetometer for attitude determination. The Sun sensor provides the angle from the spin axis to the Sun vector; the amplitude of spin signature on the Doppler data provides the instantaneous aspect angle from the spin axis to the ground station. These two measurements establish the spin axis direction. This paper describes the established theory and a method for processing the spin signature and determining the spin axis as applicable to the RBSP spacecraft. Further, RF-related errors on the Doppler carrier frequency measurement are characterized and related to the error in the spin axis orientation determination. Finally, this paper introduces the concept of attitude determination by processing the Doppler spin signature data alone over a period of time that encompasses a sufficient change to the spin axis/ground station aspect angle. This provides a further level of redundancy in the event of a failed Sun sensor. Results from processing actual data from a current mission are given.