Sun Sensor Research Articles

Autonomous Navigation Method for Mars Rover Using Distributed EKF-SLAM Assisted by Sun Sensor

Since the existing SLAM algorithm can not meet the actual demand of the Mars rover autonomous navigation problem in real-time and accuracy, a distributed EKF-SLAM algorithm is proposed based on heading assistance to achieve the rover's autonomous navigation. The solar azimuth is got by using the dual axis analog sun sensor, and then the rover heading information is calculated and added it to each subsystem of SLAM. Consequently, a SLAM model of EKF-SLAM system is built and the federal EKF is adopted to realize the state estimation of the distributed SLAM system. Finally, a whole astronomical heading assistant distributed system is constructed. The experiment test is performed in outdoor experimental environment using a mobile robot, and the experimental results demonstrate the accuracy and effectiveness of the proposed algorithm.

Low work-function tether Deorbit Kit

This works presents a system level analysis of a Deorbit Kit (DK) based on electrodynamic tether technology. The analysis is focused on two relevant scenarios for deorbiting space debris: (i) Earth Observation (EO) satellites with mass in the range of 700 kg -1000 kg and initial orbital altitude of 800 km and 98° inclination, and (ii) Mega Constellation (MC) spacecraft in the order of 200 kg and initial orbit at 1200 km of altitude and 90° of inclination. The scenarios have been selected considering the orbits that are already suffering from the space debris problem or will suffer in the next future. The DK implements a bare electrodynamic tether for capturing electrons passively from the ambient plasma while three different methods are considered for emitting the electrons back to the plasma to reach a steady electrical current on the tether. The three studied options to close the electrical circuit are: (a) a hollow cathode, which has a high technological maturity but needs expellant and a little of power, (b) a thermionic emitter, which does not involve expellant but needs power, and (c) a Low Work-function Tether (LWT) that does not need neither expellant nor power because it has a segment coated with a special material that emits electrons passively through the thermionic and photoelectric effects. In order to provide a fully autonomous operation even in case of critical failure of the mother spacecraft, the DK includes a deployment mechanism, a telemetry and telecommand system, a complete Attitude Determination and Control System with attitude sensors (GNSS, sun sensors, magnetometer) and actuators (magneto torquers), solar panels and batteries. Upon activation, the DK autonomously de-tumbles the satellite, deploys a tether and carries out the satellite's de-orbiting. The study presents DK architectures, mass budgets and simulation results for the two scenarios. It is shown that a complete DK with mass below 6% the mass of the host spacecraft can deorbit EO and MC satellites in about 1.5 years and 10 years, respectively. The importance of the development of the LWT concept to enhance the simplicity and reduce the mass, power and volume budget is highlighted.

Attitude determination by unscented Kalman filter and solar panels as sun sensor

CubeSats have became a common platform for education and technology demonstration. To overcome restrictions on CubeSat design, the Attitude Determination and Control System can use the photovoltaic panels to replace sun sensor. This work presents an attitude estimation method based on an unscented Kalman filter, using low cost gyroscope and magnetometer sensors and solar panels as a sun sensor. Data are simulated considering the influence of the Earth albedo on solar panels, and results are compared to those obtained by an extended Kalman filter.

Removing Rate Unobservability in Sun-Heading Filters Without Rate Gyroscopes

In a sun-heading determination scenario, coarse sun sensors (CSSs) can be paired with rate gyroscopes in order to estimate sun-line direction and rotation rate. Relying solely on CSS measurements for sun-heading and spacecraft rotation rate estimations is advantageous in scenarios where reliance on the fewest number of devices is desired. Here, the challenge is to find a robust method for heading determination relying neither on rate gyroscopes nor on spacecraft dynamics. In such a scenario, the rotation rate of the spacecraft is estimated in order to provide state derivative control or simply for better sun-heading estimation. A novel sun-heading filter is derived that estimates only the observable components of the body-rate vector as the rate about the sun-heading axis remains unobservable. Elegantly switching between kinematic formulations to avoid singularities of one description, it provides not only better sun-heading estimates but also a partial body-rate estimate. The new filter is compared to two filters for gyroscopes-less sun-heading estimation. One comparison filter uses a projection method to remove the unobservable rate component, and another comparison filter uses numerical heading differences to estimate a rotation rate. The filters vary in state vectors, kinematics, and filter types, with the goal of controlling or removing nonobservability.

Real-time implementation of a robust simplified intelligent proportional–integral control for CubeSat attitude determination system

This paper presents a real-time implementation of a novel simplified intelligent (SI PI) control for a low-cost attitude determination and control system (ADCS) for CubeSat. As small satellite missions are increasing, the CubeSat requires precise ADCS with attitude drift adjustment. This attitude drift if not properly compensated, the attitude knowledge will be lost as the error will increase between the actual and estimated attitudes. The proposed ADCS comprises two steps; the attitude determination which estimates the current CubeSat’s attitude and a novel simplified intelligent proportional-integral (SI PI) control algorithm that accurately adjusts the attitude. The control algorithm has no controller gains parameters and is based on the multi degree-of-freedom concept. To correct the attitude drift, the proposed ADCS utilizes magnetometer, sun sensor, and a micro-electro-mechanical (MEMS) gyroscope sensor which offers a comparative attitude that is utilized to update the estimated attitude delivered to the Kalman filter to determine the CubeSat’s angular velocity and attitude. ADCS model validation and verification are performed via MATLAB R2019b and hardware implementation. Comparison with other ADCS techniques is presented. ADCS simulated model proves precision results with error no more than 0.1 degree.

Initial Attitude Acquisition in-orbit Status of ALSAT-2A Remote Sensing Satellite

- In this paper we present the in orbit performance during the initial attitude acquisition, immediately after separation from the final stage of the launcher, until the satellite converges toward a sun pointing. On 12th july 2010 ALSAT-2A microsatellite was launched into a 670 km sun synchronous orbit, with a solar local time at an ascending node of 22h15. In the initial acquisition mode or even in the safe mode, only sun sensors and magnetometer are used for attitude determination. Knowing that, the satellite once is separated from the launcher it starts tumbling. So, as to detumble the satellite, a strategy consisting of three phases is set. The first phase consists on reducing the velocity using only the magnetorquers, until reaching the threshold angular momentum of 0.05 Nms, once this is done, the second phase is automatically enabled in such a way, the four reaction wheels are ON and speed up to get an angular momentum of -0.15 Nms along the satellite x-axis. At the end of this phase the satellite is completely detumbled and a third phase is carried out, in which the satellite minus X axis is pointed toward sun with a small rotation.

Attitude estimation and magnetometer calibration using reconfigurable TRIAD+filtering approach

This paper proposes using TRIAD and Unscented Kalman Filter (UKF) algorithms in a sequential architecture as a part of a small satellite attitude estimation algorithm. This TRIAD+UKF approach can both provide accurate attitude estimates for the satellite and calibrate the magnetometers in real-time. A complete calibration model for the magnetometers, considering bias, scale factor, soft iron and nonorthogonality errors, is assumed. In the algorithm's first stage, the TRIAD uses the available vector measurements (e.g. from Sun sensor and magnetometers) to estimate a coarse attitude. In the second stage, these estimated values are used as quaternion measurements in the UKF algorithm, together with the angular rate measured by a triad of gyros. The state vector for the UKF is composed of the attitude parameters, gyro biases and the magnetometer error terms. In result, we get fine attitude estimate for the satellite by calibrating the magnetometers in real-time. Rather high computational load of the algorithm, mainly due to the additional states, is decreased by filter reconfiguration. We evaluate the algorithm for a hypothetical nanosatellite by numerical simulations. The results show that the attitude of the satellite can be estimated with accuracy better than 0.5deg and the magnetometers can be fully calibrated.

A Hardware Implementation of Flexible Attitude Determination and Control System for Two-Axis-Stabilized CubeSat

This paper presents a hardware implementation of flexible and low-cost attitude determination and control system (ADCS) for two-axis-stabilized CubeSat. As small satellite missions are increasing, the CubeSat requires precise ADCS with attitude drift adjustment. This attitude drift if not properly compensated, will cause a slow attitude information loss as the error in attitude rises between the actual and estimated signals. The proposed ADCS comprises two steps; the attitude determination which estimates the current CubeSat’s attitude and a novel simplified intelligent proportional-integral control algorithm that accurately adjusts the attitude. The control algorithm is based on the multi degree-of-freedom controller concept and has no controller gains parameters. The proposed ADCS employs sun sensor, magnetometer, and a micro-electro-mechanical gyroscope sensor to correct the attitude drift by offering a comparative attitude that is utilized for updating the estimated attitude delivered to the Kalman filter for determining the CubeSat’s attitude and angular velocity. The ADCS model verification and validation are accomplished via Matlab/Simulink and hardware implementation. A comparison with other ADCS techniques is presented. The ADCS simulated model demonstrates precision results with error of less than 0.1°.

Design and analysis of attitude control algorithm for low earth orbiting satellite with magnetic torquer concepts using nonlinear unscented Kalman filter

Attitude control system plays the important role for to maintain the satellite to desired orientation. To control the satellite it is necessary to do the attitude stabilization. Attitude stabilization achieved by Star sensor, sun sensor, Earth sensors. Attitude control is mainly use for antenna pointing accuracy, camera focus to earth surface and solar panel pointing toward sun. Due to tumbling effects, satellite will rotate all the direction in the space. To maintain the orientation of the satellite it is necessary to design the attitude determination and control. Satellite consider as the rigid body. Inertia matrix describes the rigid body dynamics. The orientation of the satellite determine by Euler angle and Quaternion. The Low earth orbiting satellite will have enormous amount of aerodynamic drag stinking the satellite body and gravitational attraction another problem. Because of that satellite dwell, time is reduced. It means satellite more time spending particular part of the earth. The attitude estimation is measures by the orientation of vectors. Attitude estimation means to find the position and orientation of flying object with respect to the fixed reference of reference. Vector remains considered in the frame of reference to compute for find the orientation of the body of the satellite in the inertial reference system. The Earth is an inertial reference frame, and Satellite is a body frame. Attitude sensor used to measure the satellite orientation in the reference frame. This will help in accurately predicting the orbit deviation and a control system to correct if any by providing the satellite momentum means ‘mass in motion’ changes in a body rapidly in Low earth orbit due to centripetal force acting on a satellite. Attitude control system (ACS) need the numerical simulation to find the required torque demand by the help of difference between reference input (Attitude) signal and feedback signal measure by attitude sensor to trim the control surface maintain the actuator required orientation . The results will consist of two parts the first part consisting of the attitude estimation using Euler angle and Quaternion method, second part consist of estimate the control torque from magnetic torquer and error estimation using non-linear filter (Unscented Kalman Filter) with MATLAB simulation. Â

High-precision CubeSat sun sensor coupled with infrared Earth horizon detector

This paper presents a novel miniature module of orientation sensors for CubeSat nanosatellites. A high-precision sun sensor and a detector of the Earth horizon are located in one small module body. A two-axis digital sun sensor built into the orientation module is based on the registration of the sunspot formed on the matrix CMOS image sensor when the sun’s rays pass through a pinhole in the front mask. A commercially available miniature array of infrared thermometers is used as the Earth horizon sensor. The paper is mainly devoted to the development of the sun sensor. For laboratory tests, a prototype of the orientation module was developed. Laboratory tests and sun sensor calibration confirmed the sun sensor field of view of 120° and made it possible to determine its angular resolution of ±0.015°. The final accuracy of determining the direction to the Sun by the sensor that was calibrated using the existing experimental setup is ±0.15° with 95% confidence. It is possible to improve the accuracy of the sun sensor by improving the experimental setup for its calibration.

Implementation and Realisation of Micro Digital Sun Sensor Processing in FPGA

AbstractThis paper presents the implementation and realization of complex processing algorithms of newly developed, high accuracy, compact, two axes Micro Digital Sun Sensor (MDSS) using linear CMOS detector and N‐shaped slit mask for Indian Satellite missions. Sun sensor provides the orientation of satellite with respect to Sun. It is one of the prime attitude sensors for spacecraft. Due to low mass, compact size and low power consumption, the MDSS will be used for future micro/nano‐satellite missions as well as for communication Satellite missions. The electrical model of MDSS is built in Laboratory for Electro‐Optics Systems (LEOS). It is around 10 times lesser in weight and 8 times compact than the high accuracy Sun sensors presently available in ISRO. The MDSS has accuracy of 0.02°. In this paper, we have presented details of the simulation and test results of Micro DSS processing.The MDSS consists of an optical head and FPGA based processing unit. The optical head consists of CMOS linear array detector and N shaped slit mask. The N slit mask is mounted at a predefined height above the detector to provide the sensor field of view of ±64° x ±64°. The sunlight falls on the detector through the N‐shaped slit and three Sun spots are formed on the detector. After removing the unwanted signals from the images using auto threshold method, the image identification and centroid computation are carried out using moment method algorithm in real time and using the real time image clustering and centroid computing algorithm. By applying sensor transfer function, the 3 centroid information are converted to Sun aspect angle information which are the sensor outputs used for determining orientation of Satellite.All the processing is done in a single FPGA, like control signals generation for detector, real time centroid calculations, complex image clustering, transfer equation realization, interfacing with output, command module and MIL‐STD‐1553B communication. The complex transfer equation which includes trigonometric functions is implemented inside the FPGA using Co‐Ordinate Rotation Digital Computer (CORDIC) algorithm without any processor. The processing unit consists of ACTEL RTAX or Proasic FPGA with MIL‐STD 1553B and LVDS peripheral interface.

Hybrid loop controlled dual axis mechanical solar tracking systems and MPPTfor Nano/Micro grid Applications

This paper emphasizes the important role of solar tracker systems in Micro/Nano grid level applications. Basically mechanical and electrical type trackers are different in operation but they are similar in way of tracking of solar energy. Mechanical trackers (Dual axis solar tracking system) are widely used where space is less and quality is more and with fixed panels solar MPPT is adoptable. Hybrid loop control which controls the relative motion of solar tracking system (STS) during the day as per the position of sun. Hybrid loop control (HLC) is effective rather than open and closed loop systems due its mode of control using sun position algorithm and sensors based as per the requirement and environmental conditions. Most commonly used sensor based trackers takes feedback to adjust the panel position in such a way to track the sun, but the situation when the sky is cloudy creates problem in performance of the system. The estimated accuracy is very close to ±0.0003º but it depends on the irradiation coming from sun. So, Sun tracking system by using solar position algorithm with accuracy is required in mechanical solar trackers and maximum power point tracking in electrical trackers. In order improve the efficiency of whole solar tracking system, we incorporated both open and closed loops i.e., hybrid loop control which plays important role to improve the overall performance. In view of that we incorporated both open and closed loops i.e., hybrid loop control. Here we presented Maximum power point tracking (MPPT) which is electrical tracking system simulated in MATLAB environment and also presented mechanical type of tracking system It is facilitated by using Programmable logic controller (PLC) to access more input/output controls and simulation has been done in CODESYS software using ladder logic programming.

A test and calibration environment for the Brazilian star tracker

For many spacecraft to accomplish their mission goals there is a need to have an accurate control of their spatial orientation (attitude), so that instruments aboard the spacecraft, such as remote sensing cameras, telescopes, and other scientific payload be correctly pointed to their targets. For that purpose, many spacecraft employ sophisticated attitude control systems, fed by a variety of attitude sensors: sun sensors, horizon sensors, gyroscopes, star trackers, etc. Among the attitude sensors capable of providing an absolute attitude measurement, star trackers are considered the most accurate. This work presents an overview of the test and calibration infrastructure built for an autonomous star tracker (AST) being developed by the Electro-Optics group of INPE, result of an effort to increase the competence of the country in attitude control systems.

Efficient hyper-parameter determination for regularised linear BRDF parameter retrieval

ABSTRACTLinear kernel driven models of the surface Bidirectional Reflectance Distribution Function (BRDF) are valuable tools for exploiting Earth observation data acquired at different sun–sensor geometries. Here we present a method that efficiently determines linear BRDF model weights using Tikhonov smoothing where the smoothing parameter λ is determined via a Generalized Singular Value Decomposition with the root mean square error prescribed depending on the MODIS band. We applied this method to twenty-six different deciduous broadleaf sites across an entire year using the MODIS Terra and Aqua reflectance data products. Kernel weights and white sky albedo derived from this GSVD method were generally consistent with those provided by the MCD43 data products. The GSVD derived results had less sample variability compared to the MCD43 data products, attributable to the assumed smoothness between kernel weights in the Tikhonov smoothing method. The GSVD technique consistently outperforms MCD43 in the reconstruction of observed MODIS reflectance data, of which retrievals from this method will do a better job of estimating albedo and normalizing data to specified geometries.

Design of on-board calibration methods for a digital sun sensor based on Levenberg–Marquardt algorithm and Kalman filters

Digital sun sensor is one of the most important sensors used in the Attitude Determination System (ADS) of the satellite. Due to the harsh environmental conditions that exist in the space, various distortions may occur in the sun sensor optical system that lead to the reduced accuracy of this equipment. So, it is necessary to recalibrate the optical parameters of the aforementioned sensors. For this purpose, first a novel attitude independent error model is proposed for the SS-411 sun sensor that includes the central point of the CCD array, installation error, filter thickness and sensor misalignment. So, the mutual interfaces between the sensor parameters are considered in the developed model. In order to extract the sensor parameters, a nonlinear optimization technique called the Levenberg–Marquardt is applied to the developed model as a batch algorithm. In addition, the Extended Kalman Filter (EKF) and the Unscented Kalman Filter (UKF) have been utilized as sequential strategies. It will be shown that by considering a worst case of variation amount for sensor parameters, an accuracy improvement of about 17° is achieved by the developed calibration algorithms. Comparison between the developed algorithms represents that UKF has higher accuracy, shorter time convergence but higher computational load.

Performance Analysis of GNSS-Aided Inertial Navigation Systems on Spinning Flight Vehicles

Inertial navigation on spin-stabilized sounding rockets and missiles is challenging due to the high spin rate about the longitudinal axis. Especially the scale factor error of the roll axis gyroscope causes an accumulating roll angle error, which may subsequently transform to the lateral axes if, for example, the rocket is actively reoriented. This orientation error leads to growing position and velocity estimation errors whenever propulsion or aerodynamic forces act on the rocket. Specialized inertial platforms were developed in the 1970s with gimbaled inertial measurement units isolating the sensors from the high spin rate. Because these platforms are rarely available on the market today, ongoing research focuses on integrated navigation systems with strap-down inertial measurement units and additional aiding sensors such as satellite navigation receivers, Sun sensors, or star tracker. Because satellite navigation receivers are well-established on sounding rockets for range safety anyway, this paper investigates to what extent position, velocity, and particularly orientation errors can be estimated with satellite navigation aiding only. Using the example of SHEFEX-2, the paper discusses the observability of translational and rotational navigation errors by translational aiding measurements along a typical trajectory of a two-staged and actively reoriented sounding rocket both qualitatively and quantitatively.

Nanosatellite attitude estimation using Kalman-type filters with non-Gaussian noise

In order to control the orientation of a satellite, it is important to estimate the attitude accurately. Time series estimation is especially important in micro and nanosatellites, whose sensors are usually low-cost and have higher noise levels than high end sensors. Also, the algorithms should be able to run on systems with very restricted computer power. In this work, we evaluate five Kalman-type filtering algorithms for attitude estimation with 3-axis magnetometer and sun sensor measurements. The Kalman-type filters are selected so that each of them is designed to mitigate one error source for the unscented Kalman filter that is used as baseline. We investigate the distribution of the magnetometer noises and show that the Student's t-distribution is a better model for them than the Gaussian distribution. We consider filter responses in four operation modes: steady state, recovery from incorrect initial state, short-term sensor noise increment, and long-term increment. We find that a Kalman-type filter designed for Student's t sensor noises has the best combination of accuracy and computational speed for these problems, which leads to a conclusion that one can achieve more improvements in estimation accuracy by using a filter that can work with heavy tailed noise than by using a nonlinearity minimizing filter that assumes Gaussian noise.

Preliminary Design of Imaging Microsatellite for Preventing Illegal Fishing in Indonesia

A constellation of imaging microsatellite system using Synthetic Aperture Radar (SAR) and Automatic Identification System (AIS) has been developed to prevent illegal fishing in Indonesia. The constellation of microsatellite is shown to be able to cover the entire area of Indonesia during 24 hours duration. Each satellite is capable of taking pictures with minimum resolution of 3 meter per pixel and swath width of 28.55 km. AIS equipped in each satellite will collect data of the ships operating on sea area of Indonesia. The information about the ships collected by AIS will then be compared to the picture captured by SAR to determine if the ships are fishing illegally or not. To support the mission, each satellite is equipped with sun sensors, earth sensors, and GPS to precisely determine its attitude and position. In addition, each satellite also equipped with Control Moment Gyros (CMGs) capable of precision pointing and high slew rate maneuver with low power consumption. The payload communication of the satellite will use VHF and X-band while the TT&C will use S-band. Furthermore, each satellite requires a total of 838 W to operate which correspond to 6 m2 of solar panel and 8.8 kg and 590 Wh of batteries. Finally, it is concluded that each satellite has 95 kg of total mass and costs around $ 5.4 million to develop.

Estimating Sun Vector Based on Limited In-Orbit Data

This paper introduces an attitude and orbit determination system, in particular a sun vector determination algorithm, for HORYU-IV nano-satellite by Kyushu Institute of Technology (Kyutech). Usually, nano-satellites cannot use sensors with high specs because of limited budget and onboard space. In order to achieve the required accuracy of attitude determination, estimation algorithms combining different sources of attitude information should be used. One of the HORYU-IV missions has required sun vector information. In order to detect the sun angle, six sun sensors, which have originally been designed and developed by the authors, and five solar panels have been used. However, due to the narrow field of view and to the saturation property of these sensors, the duration of sun detection is quite limited. At a non-detection period, a gyro sensor is used in order to propagate the attitude. However, the sensor has the bias error in its measured value. Hence, the Extended Kalman Filter has been used in order to estimate the sun vector and the bias error of gyro sensors. A simulation result shows that the algorithm can estimate the sun vector in 1deg accuracy.

A Sun Sensor Based on Regular Pyramid Sensor Arrays

Non-planar sun sensors determine solar position by existing photodiodes or solar panels without increasing the size or mass of aerospace equipment such as satellites. However, it suffers from the coarse orientation accuracy. A sun sensor that is formed by photodiodes or solar panels mounted on the lateral surfaces of regular pyramid was proposed to reduce the estimated error of the sun vector, which indirectly improve the accuracy of solar orientation determination. The orientation formulation of the sun sensor and its sun vector error were derived to evaluate the interference suppression for the sun sensor. Experimental data validates the maximum error of orientation is less than 1° in a clear sky by the sun sensor consisting of 16 solar panels, which are mounted on the lateral surfaces of a regular 16-pyramid with the output error less than ±5% and the assembly error less than ±1°.