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In this paper, the satellite attitude control system subject to parametric perturbations, external disturbances, time-varying input delays, actuator faults and saturation is studied. In order to make the controller architecture simple and practical, the closed-loop system is transformed into a disturbance-free nominal system and an equivalent disturbance firstly. The equivalent disturbance represents all above uncertainties and actuator failures of the original system. Then a robust controller is proposed in a simple composition consisting of a nominal controller and a robust compensator. The nominal controller is designed for the transformed nominal system. The robust compensator is developed from a second-order filter to restrict the influence of the equivalent disturbance. Stability analysis indicates that both attitude tracking errors and compensator states can converge into the given neighborhood of the origin in finite time. To verify the effectiveness of the proposed control law, numerical simulations are carried out in different cases. Presented results demonstrate that the high-precision attitude tracking control can be achieved by the proposed fault-tolerant control law. Furthermore, multiple system performances including the control accuracy and energy consumption index are fully discussed under a series of compensator parameters.

A pulsed electric thruster named Pulsed Electrogasdynamic (PEG) thruster was investigated. In this thruster, the propellant gas is accelerated gasdynamically to high exhaust velocities (10–15 km/s...

This paper investigates the performance of adaptive neural networks through simulations for satellite systems involving three-axis attitude control algorithms. PID tuning is the method employed traditionally. An optimally tuned, to minimizes the deviation from set point. It also responds quickly to the disturbances with some minimal overshoot. However, the disadvantage of poor performance has been observed in these controllers when manual tuning is used which in itself a monotonous process is. The PID controller using Ziegler-Nichols has more transient responses of satellite such as Overshoot, Settling time, and Steady state errors. For overcome this technique, the proposed analysis implemented an Adaptive Neural Network with PID tuning. The paper aims to combine two feedback methods by using neural networks. These methods are feed- forward and error feedback adaptive control. The research work is expected to reveal the inside working of these neural network controllers for state and error feedback input states. An error driven adaptive control systems is produced, when the neural networks acquire the knowledge of slopes and gains regarding the error feedback, while, with state feedback the system will keep trying to approximate a stable approach in order to stabilize the attitude of the satellite.

An optimal thruster configuration for attitude control subsystem of a spacecraft is presented in this paper. The optimal configuration is designed according to minimum number of required thrusters for satisfying desired reliability with specific redundancy level. The genetic algorithm is employed for optimization process and feasibility of the results is evaluated using algebraic and geometry methods. The main feature of the proposed configuration among feasible configuration with minimum number of required thrusters, which has held to optimal configuration, is that this configuration has maximum reliability and minimum fuel consumption. In addition to feasibility, attitude control performance of some configurations is also examined through the simulation. The results of simulation confirm that the proposed configuration has desirable performance. It is noteworthy to mention that the configuration with maximum number of required thrusters, which is a conventional configuration such that each thruster belongs to only one control channel, has less fuel consumption than optimal configuration. However, the total mass of optimal configuration is less than that of conventional configuration due to a smaller number of thrusters.

An active fault-tolerant satellite attitude control scheme based on fault effect classification is presented at the occurrence of faults associated with torques. In this paper, the flexibility and practicability of the fault-tolerant scheme are top priorities. Faults are modeled and divided into additive and multiplicative ones in order to estimate and deal with them specifically and exactly. The additive faults, including additive part of flywheel faults and other uncertain fault torques, are estimated by additive fault estimator and compensated on the basis of nominal controller, whereas the multiplicative faults, denoting torque gain parameter faults of flywheels, are estimated by multiplicative fault estimator and the estimated fault parameters are used for dynamic torque command distribution of flywheels. The final simulation examples and performance comparison of three fault-tolerant schemes show that the proposed scheme based on fault effect classification is an effective, flexible and saving-energy fault-tolerant satellite attitude control scheme. It possesses an engineering value for improving reliability and prolonging on-orbit working lifetime of satellites.

This paper studies the attitude control of a satellite in three-axis using only four thrusters. The mathematical model of the attitude is represented as a switched system with 16 subsystems. Each subsystem is defined based on the thruster's on/off status. The dynamic programming is utilised to find the optimal switching between subsystems such that a cost function is optimised. Furthermore, to generalise the solution for a specific domain of interest, neural network is employed for offline training and approximating the cost function. An offline training algorithm is suggested to find the optimal weights of the neurons and determine optimal switching. It is shown that the proposed method can execute a manoeuvre within fixed final time. Moreover, the control is robust against uncertainties in the system modelling. Finally, the system modelling and control approach is suggested as a framework to design low-cost attitude control system unit, suitable for microsatellite-class.

In recent years, the maturation of small satellite technology has led to their adoption for a variety of space missions. The next generation of small satellite missions, however, will likely have the satellites operating in formations or “constellations” to perform missions that are not currently possible. A key enabling technology for constellation-based missions is a miniaturized propulsion system that is capable of delivering the extremely low impulse levels required for maintaining precise relative position and orientation. Existing propulsion solutions for this regime suffer from compromises on power, safety, and cost that have limited their adoption. In this work, we describe a new, low-power micropropulsion concept based on the thermal decomposition of an inert chemical blowing agent (CBA) as the propellant. A meso-scale prototype device is designed, fabricated, and tested. The experimental results indicate that this concept, when appropriately scaled, is capable of providing thrust levels (∼1 μ N) and impulse-bits (∼0.1 μ N·s) that are commensurate with the intended application.

Solar sailing is a new satellite propulsion technology using radiation pressure exerted by sunlight on a large mirrored surface. Since it does not need propellants, it is increasingly being considered by both the European Space Agency and the National Aeronautics and Space Administration for future science missions. An attitude control system is essential for a sail craft to maintain a desired orientation. IKAROS, launched in 2010, practically proved the possibility of using a solar sail as a propulsion system. However, it also showed the current sail orientation system could change the attitude very slowly, about 1 degree per day. In contrast to the existing single solar sail design, a new distributed four-sail configuration is proposed in this paper and the coordinated motion of the four sails is used to control the attitude pointing of a satellite. The feasibility and efficiency of this proposed design were assessed and concluded that it is possible to steer a CubeSat up to 1 degree in 60 seconds for either the roll or pitch axes.

Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly, Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnológico de Aeronáutica, in São José dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnológico de Aeronáutica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel desaturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control.

This article has discussed the development of a three-axis attitude digital controller for an artificial satellite using a digital signal processor. The main motivation of this study is the attitude control system of the satellite Multi-Mission Platform, developed by the Brazilian National Institute for Space Research for application in different sort of missions. The controller design was based on the theory of the Linear Quadratic Gaussian Regulator, synthesized from the linearized model of the motion of the satellite, i.e., the kinematics and dynamics of attitude. The attitude actuators considered in this study are pairs of cold gas jets powered by a pulse width/pulse frequency modulator. In the first stage of the project development, a system controller for continuous time was studied with the aim of testing the adequacy of the adopted control. The next steps had included an analysis of discretization techniques, the setting time of sampling rate, and the testing of the digital version of the Linear Quadratic Gaussian Regulator controller in the MATLAB/SIMULINK. To fulfill the study, the controller was implemented in a digital signal processor, specifically the Blackfin BF537 from Analog Devices, along with the pulse width/pulse frequency modulator. The validation tests used a scheme of co-simulation, where the model of the satellite was simulated in MATLAB/SIMULINK, while the controller and modulator were processed in the digital signal processor with a tool called Processor-In-the-Loop, which acted as a data communication link between both environments.function and required time to achieve a given mission accuracy are determined, and results are provided as illustration.

This paper extends the previous works that appeared in Acta Astronautica. An approach that incorporates the Active Force Control (AFC) technique into a conventional Proportional-Derivative (PD) controller is proposed for a 50 kg small satellite. Numerical treatments are performed to validate the effectiveness of AFC. The attitude control capability of the combined energy and attitude control system (CEACS) is expected to improve. The result shows an important attitude pointing enhancement for the CEACS attitude control task.

A low-mass, low-volume propulsion subsystems based on electrically activated micro-thrusters that utilize chemically-inert solid propellants are beneficial for small satellite attitude control applications. Micro-thrusters are able to deliver small impulse bits of about several µNs to satellites and characterized by simplicity, scalability, low cost, low weight and high reliability. In this work measurements of the key parameters of micro-Cathode Arc Thruster (µCAT) are presented and µCAT performance compared with other commercially available thruster technologies operating at the same power range.

model reduction methods, it is possible to reduce the controller order to a lower level with acceptabledegradation in control performance. However, using controller approximationwith fewer states implies an increase in 1. Still, there are open numerical problems in connectionwith actual implementation in 1-synthesis. 2) The robust performance (RP) condition in Eq. (8) is identical to a robust stability (RS) condition with an additional perturbation block 1P . In other words, the RP condition is more severe than the RS condition, and so there is no doubt of RS in this illustrative satellite case, where the RP condition is satisŽ ed.

The paper proposes the use of a bitethered pendulum mass deployed from satellite for open-loop satellite attitude control in elliptic orbits. A simple approximate analytical approach has been adopted to develop control law governing moments generated through a judicious non-feedback scheme of tether length variation with a view to neutralize the adverse effect of the harmonic excitation induced by eccentricity. The detailed simulation of the system response shows that the proposed control strategy enables much higher pointing accuracies for satellites in elliptic orbits, even with most adverse satellite mass distribution. The small auxiliary pendulum mass needed along with modest tether lengths make the concept particularly attractive.

Following several flight and ground test failures of spacecraft systems using single-shot, 'normally closed' pyrotechnically actuated valves (pyrovalves), a Government/Industry cooperative program was initiated to assess the functional performance of five qualified designs. The goal of the program was to provide information on functional performance of pyrovalves to allow users the opportunity to improve procurement requirements. Specific objectives included the demonstration of performance test methods, the seating; these gases/particles entered the fluid path of measurement of 'blowby' (the passage of gases from the pyrotechnic energy source around the activating piston into the valve's fluid path), and the quantification of functional margins for each design. Experiments were conducted at NASA's Langley Research Center on several units for each of the five valve designs. The test methods used for this program measured the forces and energies required to actuate the valves, as well as the energies and the pressures (where possible) delivered by the pyrotechnic sources. Functional performance ranged widely among the designs. Blowby cannot be prevented by o-ring seals; metal-to-metal seals were effective. Functional margin was determined by dividing the energy delivered by the pyrotechnic sources in excess to that required to accomplish the function by the energy required for that function. Two of the five designs had inadequate functional margins with the pyrotechnic cartridges evaluated.

The paper describes ECUMES, an approach to a space based ocean altimetry mission configured by Matra Marconi and Fairchild to perform a follow-on of TOPEX/Poseidon.

A safehold mode has been implemented for Hubble Space Telescope that ensures a power-positive and thermally safe state without the use of rate gyros. This paper presents an overview of this zero-gyro sunpoint safemode, followed by details of the algorithm, the control system design, and results of simulations and of an on-orbit test of the algorithm.

Part I of this work deals with the use of electrodynamic forces for control of tethered satellite system. A system formed by two massive end-bodies connected to each other by a current carrying tether is to be kept in an Earth-pointing orientation by means of joint actions of thrusters on one of the end-bodies and electrodynamic forces acting on the tether. >

Results on star-mapper data for attitude determination based on observations carried out over fifteen reference great circle's (RGC) are presented and discussed. Precision on the three-angle attitude reconstruction is of the order of 60 milli-arcsec along-scan and 100 and 180 milli-arcsec normal-to-scan. With this first year data, it is possible to correct the star position from the residuals of the attitude solution. Satisfactory corrections are obtained for RGC's abscissa and used at a latter stage of the reduction to remove grid-step ambiguity in the RGC reduction.

This paper deals with torque command generation using single gimbal control moment gyros. The angular momentum and torque envelopes are assumed to be known a priori. A method based on back integration of the gyro torque equation from desired final conditions is used to determine a family of initial gimbal angles that avoid singularities. Each member of this family is defined as a preferred initial gimbal angle set. The pseudoin- verse steering law is used during the numerical integrations. This procedure is demonstrated by means of numerical examples that include attitude control and momentum management of the Space Station Freedom. A feedback control scheme based on "null motion" is also developed to position the gimbals at preferred angles. ONTROL moment gyros (CMGs) are attractive space- craft attitude-control devices. They require no expendable propellant, which is a limited resource and can contaminate the spacecraft environment. Their fixed rotor speeds minimize structural dynamic excitations. They can be used for rapid slewing maneuvers and precision pointing. For low Earth or- biting spacecraft, momentum dumping can be easily achieved by gravity-gradient torques. From the steering-law viewpoint, it is widely accepted that double-gimbal CMGs (DCMGs) are preferable to single-gimbal CMGs (SCMGs). For DCMGs, steering laws proposed by Kennel 1'2 have been well accepted. The SCMGs have the advantages of possessing relative me- chanical simplicity and producing amplified torques (for low spacecraft angular velocities) on the spacecraft. However, de- velopment of gimbal steering laws for their use is made diffi- cult by the existence of internal singular states. For any system of n CMGs and any direction in space, there exist 2" sets of gimbal angles3 for which no torque can be produced in that direction, and these sets are called internal singularities. Exter- nal singular states correspond to directional angular momen- tum saturation. DCMGs have internal singularities also, but they are easier to avoid. Margulies and Aubrun3 present a geometric theory of SCMG systems. They characterize the momentum envelope of a cluster of SCMGs and identify the internal singular states. Yoshikawa4 presents a steering law for a roof-type configura- tion with four SCMGs. His steering law is based on making all of the internal singular states unstable by providing two jumps with hystereses around the singularities. Cornick5 develops singularity avoidance control laws for the pyramid configura- tion. His technique is based on the ability to calculate the instantaneous locations of all singularities. Hefner and McKenzie6 develop a technique for maximizing the minimum torque capability of a cluster of SCMGs in the pyramid con- figuration. Bauer7 concludes that it is impossible to avoid some singularities and, in general, no global singularity-avoid- ance steering law can exist. Consequently, there will be in- stances when torque demand cannot be met exactly.

A new and straightforward approach to mixed psynthesis is proposed. The main idea is that the structure singular value can be equivalent to an weighted H2 system norm by shaping a frequency-dependent weight. The resulting controller obtained from a sequence of weighted H2 optimizations may possess robust performance property from p criterion. Two successful designs are given to stress the superior ability of the H2-based loop-shaping method for mixed p-synthesis.

A novel silver detector incorporated in the Long Duration Exposure Facility (LDEF) has been used to measure the attitude stability of the vehicle. It is thereby established that the LDEF spacecraft maintained a very stable attitude during its nearly 6-year-long flight, notwithstanding a small offset yaw of 8.0 +/- 0.4 deg clockwise from nominal attitude (as viewed from space) and a +/- 0.2-deg oscillation about this offset yaw. LDEF experiments which depend on orientation relative to the forward direction may need to be corrected for the angular offset.

The methods utilized to design orbiter attitudes that satisfied the various Space Shuttle program and payload constraints of the Astro-1 Spacelab astronomy mission are described. The premission methods developed for attitude design and how those methods were applied to accomplish the large amount of real-time replanning required during the flight are presented. Some of the attitude concerns for Astro-1 included orbiter window thermal constraints, two separate telescope fields-of-view, tracking and data relay satellite Ku-band access, and telescope sun avoidance.

In 1967 a paper at the AIAA Guidance, Control and Flight Dynamics Conference in Huntsville, Ala. presented for the first time the prot)osed SKYLAB Attitude and Pointing Control System (APCS) The system requirements, Apollo Telescope Mount (ATM) configuration, control philosophy, and operational modes were presented and the APCS described. The Initial mission and system design requirements changed during the period of time before the SKYLAB was launched. This paper will review the Initial and final APCS requirements and goals and their relationship. The actual flight mission (and Its alterations during the flight) and known achieved APCS performance will then be presented. SKYLAB was a tremendous success in furthering man's scientific knowledge; but perhaps SKYLAB will be remembered more for the anomalies and the efforts undertaken to solve them. On May 14, 1973, the unmanned SKYLAB Orbital Workshop (OWS) was launched from Cape Kennedy. Serious hardware failures began to occur during ascent through the atmosphere and their spectre continued to haunt both the astronauts and their ground based support team. Nor were these the only surprises affecting the design and operation of the APCS. Mission requirements for pointing to various stellar targets and to nadir for earth resources experiments were added after the hardware was designed. The chance appearance of comet Kohoutek during the SKYLAB operational life-time caused NASA to add comet observation to the mission requirements and to adjust the time when the third crew would man the SKYLAB. The development of new procedures and software for the opportunity to observe this visitor to our solar system is described.

GMS has been operated quite normally since it was launched on 14 July 1977, except for the following anomalies: PMT power source and spacecraft spin axis tilt. The influence of these anomalies on the output images are reduced to negligible value by appropriate spacecraft operation and data processing.

The accuracy of presently available IR horizon sensors is not sufficient to meet the stringent attitude sensing and control requirements for future remote sensing and meteorological satellites. The different sources of error in a horizon sensor are analyzed. The accuracy of the sensor is presently limited by the detector noise. Use of HgCdTe in place of an immersed bolometer detector, which is used in conventional horizon sensors eliminates many of the errors. Hence, it is possible to design an ultimate IR horizon sensor whose accuracy is limited only by the uncertainty of the Earth horizon. Comparison of performances of the two types of detectors for horizon sensing is given and possible configurations of sensor using this detector are discussed.

The Institute of Space and Aeronautical Science (ISAS), University of Tokyo, has been engaged in the development of scientific satellites and their launch vehicles of Mu series. ISAS has conducted the first flight test of M-3S, the newest version of Mu series, and launched a technology test satellite MS-T4 into an orbit. The objectives were to verify the overall performance of the vehicle and to perform various experiments on the satellite technology to be applied to the following scientific satellites. Technological aspects of M-3S and MS-T4 are reviewed in reference to the recent trend in the scientific satllite program.

Attitude performance of GEOS 2 gravity gradient spacecraft, discussing prelaunch design and postlaunch analyses

Single nozzle thermal storage resistojet thrustor for low thrust satellite attitude and orbit control applications, noting design and performance

This paper presents the results of a study of the dynamics of a spinning Skylab space station. The stability of motion of several simplified models with flexible appendages was investigated. A digital simulation model that more accurately portrays the complex Skylab vehicle is described, and simulation results are compared with analytically derived results.