Design Algorithm for Upper Stage’s Attitude Control System

Space Qualification is a vital component of successful spacecraft design. The entire scope of the space qualification process is extremely large. This thesis will focus primarily on the Attitude Control System and Space Environment model, with the intent of furthering the verification of the design for the Joint Australian Engineering Satellite (JAESAT). The Attitude Control System proposed for use aboard JAESAT is a combination of permanent magnets with hysteresis rod damping. This system has proved effective on many previous satellite missions, however analysis must be undertaken to optimise the control provided for the individual design proposed. The Space Environment will affect the operation of many of the subsystems aboard JAESAT. Knowledge of the magnitude of such affects is essential to prevent complications through improved design. To do this requires an accurate model of the Space Environment local to JAESAT. Several models for different environmental characteristics are presented as part of this dissertation. Further detailed modelling was performed using the Space Environment Information System, an online software package provided by the European Space Agency. Some of the many results that have been obtained from the Space Environment Information System are presented in graphical format. A detailed model of the Attitude Control System was constructed using Matlab. The environmental disturbance torques were calculated using the environmental conditions previously modelled. Several simulations were conducted yielding mixed results. Generally, it was found that a permanent magnet strength of 14 A.m2 and spin rate of 0.4 revolutions per minute is sufficient to control JAESAT to within 1° of the magnetic field vector. Numerical simulation of the hysteresis rod damping however, has been shown to be ineffective and additional testing is required to successfully implement the hysteresis effects in the system simulation. An analysis of the model revealed that there were minimal sources of error present in the simulation process. The only significant source of error was due to the numerical integration of the equations associated with the satellite motion. Another possible source of model inaccuracy was due to the variation between the numerically simulated space environment and the actual space environment. This dissimilarity however, will only result in small changes of satellite orientation and will not have any effect on the satellite attitude. In the future, the Space Environment and Attitude Control System model that was constructed should be updated to reflect any changes in satellite design and orbit parameters. Simulations of the final Attitude Control System hardware should also be conducted in order to show that adequate attitude control is maintained throughout a period of several orbits. The hysteresis rods should be tested in order to determine their damping properties, and can then be incorporated in the satellite attitude simulations. Based upon numerous simulation results and research, it was concluded that the Attitude Control System that has been proposed was sufficient to provide the required level of attitude control for JAESAT. The successful verification of the Attitude Control System operation implies that the system is Space Qualified. Exact specifications for the Attitude Control System hardware are dependant upon the satellite orbit and design requirements. The model of the space environment is also sufficiently accurate to use in the verification of other satellite systems.

A GROUND LOOP ATTITUDE CONTROL SYSTEM FOR ANIK B

The Naval Research Laboratory (NRL) has developed a new bipropellant Upper Stage providing orbit transfer propulsion capabilities from a Delta II launch vehicle. This bipropellant stage utilizes a pair of Composite Overwrapped Pressure Vessels (COPV) to store helium for pressurization of the Upper Stage fuel and oxidizer propellant tanks. A systems engineering trade study was performed to determine the size requirements of the helium tanks followed by competitive vendor selection and subsequent tank development and qualification program. This paper discusses the helium tank sizing analysis and documents details of the COPV design and qualification testing effort. Additionally, spacecraft installation interfaces, a tank pressure hazard analysis, and on-orbit temperature and pressure performance are summarized and presented. Program description. The Naval Research Laboratory (NRL) is the corporate research laboratory for the United States Navy and Marine Corps and conducts a broad program of scientific research, technology and advanced development. The NRL, sponsored by the Defense Advanced Research Projects Agency (DARPA), developed the Upper Stage (US) for the Microsatellite Technology Experiment (MiTEx) program by drawing on its corporate experience building and operating spacecraft. The Upper Stage provided the orbit transfer propulsion for the MiTEx microsatellites and was launched on a Delta II in June of 2006. NRL personnel performed a series of mission analyses and trade studies that showed that an earth storable liquid bipropellant system would best satisfy the MiTEx orbital maneuvering requirements. A constant feed pressure is required to optimize bipropellant engine performance; therefore the propulsion subsystem design includes two Composite Overwrapped Pressure Vessels (COPV) to store helium pressurant. ARDE, Inc was selected as the COPV contractor after a competitive procurement. ARDE developed a tank design to meet the NRL requirements, and conducted extensive tests both to qualify the design and to provide thermal – helium expulsion data to support NRL mission planning needs. Mission sizing and blowdown. NRL personnel performed a series of analyses to support the COPV procurement. These included parametric studies, mission requirements flowdown, budget allocations to the various components including the COPV, and packaging studies. The key requirement addressed in these analyses was the pressure remaining in the COPV at the end of each maneuver and 1 Osborn, Clauss, Gorin, Netwall, “Micro-satellite Technology Experiment (MiTEx) Upper Stage Propulsion System Development,” AIAA 2007-5434 how well that pressure could satisfy feed system flow requirements. These analyses considered the propellant tank volume, the propellant volume required for various orbital maneuvers, the propellant volumetric flow rate, and the COPV’s operating pressure at the beginning of the maneuver to develop a trade-space that included helium end of maneuver pressure and helium flow path pressure drop. The parametric analyses showed that the temperature of the helium at the end of the maneuver is the driver for the final expected pressure. The helium tank blowdown can be modeled as a polytropic process using Equation 1 where n represents a constant for a particular process.

In this paper, a health status assessment scheme is studied for the attitude control system (ACS) of fixed-wing unmanned aerial vehicle (FWUAV) based on an improved multivariate state estimation technology that incorporates a dynamic memory matrix. Firstly, the parameters of the FWUAV representing the health status of the ACS are selected as feature parameters, and the historical health data of the feature parameters for the FWUAV is compared with the real-time test data. At the same time, the multivariate state estimation technology is applied to obtain the abnormal degree of components for the ACS. Based on the analytic hierarchy process and the expert experience, the weights are obtained for the different elements at the same functional level of the ACS. Secondly, the functional structure of the ACS is analyzed, and the abnormal degree is calculated for the ACS by combining the concept of reconfigurability and the weight information of each element, and the health status assessment indicators are further established. The health status result for the ACS of the FWUAV is acquired according to the efficiency value of the indicators and the abnormal degree of the ACS. Finally, the effectiveness of the developed algorithm is verified by the simulation analysis.

To improve the reliability of the pico-satellites' attitude control system,new methods were proposed,which employed magnetic coils as the only actuators to re-capture the attitude in case of flywheel's failure.Two algorithms,a PD controller and a linear quadratic regulator,were applied to stabilize pico-satellite to nadir pointing using fully magnetic coils.When error occurs,the whole control system switched to the fully magnetic control mode to recover the attitude.Meanwhile,numerical simulations were also taken under the situation that only magnetic actuators available to the attitude control system of the pico-satellites.Results show that these two algorithms can achieve a certain accuracy of nadir-pointing in both situations.The attitude recovery accuracy adopting the PD control law better than 0.5° while the one with the LQR control law attains 10° or so.

The Engineering Systems Design Lab (ESDL) at the University of Illinois introduced Strain-Actuated Solar Arrays (SASAs) as a solution for precise satellite Attitude Control System (ACSs). SASA is designed to provide active mechanical vibration (jitter) cancellation, as well as small slew maneuver capabilities to hold a pose for short time periods. Current SASA implementations utilize piezoelectric distributed actuators to strain deployable structures, and the resulting momentum transfer rotates the spacecraft bus. A core disadvantage, however, is small strain and slew capability. Initial SASA systems could help improve pointing accuracy, but must be coupled with another ACS technology to produce large reorientations. A novel extension of the original SASA system is presented here that overcomes the small-displacement limitation, enabling use of SASA as a sole ACS for some missions, or in conjunction with other ACSs. This extension, known as Multifunctional Structures for Attitude Control (MSAC), can produce arbitrarily-large rotations, and has the potential to scale to large spacecraft. The system utilizes existing flexible deployable structures (such as solar arrays or radiators) as multifunctional devices. This multi-role use of solar panels extends their utility at a low mass penalty, while increasing reliability of the spacecraft ACS.

The problem of trajectory optimization is important in all space missions. The solution of this problem enables one to specify the optimum thrust steering program which should be followed to achieve a specified mission objective, simultaneously satisfying the constraints.1 It is well known that whether or not the ascent trajectory is optimal can have a significant impact on propellant usage for a given payload, or on payload weight for the same gross vehicle weight.2 Consequently, ascent guidance commands are usually optimized in some fashion. Multi-stage vehicles add complexity to this analysis process as changes in vehicle properties in one stage propagate to the other stages through gear ratios and changes in the optimal trajectory. These effects can cause an increase in analysis time as more variables are added and convergence of the optimizer to system closure requires more analysis iterations. In this paper, an approach to simplifying this multi-stage problem through the creation of an upper stage capability boundary is presented. This work was completed as part of a larger study focused on trade space exploration for the advanced booster system that will eventually form a part of NASA s new Space Launch System.3 The approach developed leverages Design of Experiments and Surrogate Modeling4 techniques to create a predictive model of the SLS upper stage performance. The design of the SLS core stages is considered fixed for the purposes of this study, which results in trajectory parameters such as staging conditions being the only variables relevant to the upper stage. Through the creation of a surrogate model, which takes staging conditions as inputs and predicts the payload mass delivered by the SLS upper stage to a reference orbit as the response, it is possible to identify a of staging conditions which all satisfy the SLS requirement of placing 130 metric tons into low-Earth orbit (LEO).3 This identified surface represents the 130 metric ton capability boundary for the upper stage, such that if the combined first stage and boosters can achieve any one staging point on that surface, then the design is identified as feasible. With the surrogate model created, design and analysis of advanced booster concepts is streamlined, as optimization of the upper stage trajectory is no longer required in every design loop.

During an impulsive orbital maneuver, thrust vector misalignment from the center of mass (C.M) generates a large exogenous disturbance torque that results in attitude deviation. This paper aims not to use the reaction control system (RCS) for the spacecraft attitude control. In order to reject the large disturbance very fast, a new control system is proposed. In this method, the large disturbance torque is rejected quickly while the RCS is not employed. The control system is based on one degree of freedom (1DoF) gimbaled-thruster, spin-stabilization, and two control moment gyros (CMG). The nonlinear two-body dynamics of the mentioned spacecraft is formulated. Because RCS is not used, this method is an efficient and implementable method for attitude control of small spacecraft. Numerical simulation shows that thrust vector deviation converges to zero despite disturbance torques. By this method, the disturbance is rejected very fast; thus an accurate orbital velocity change can be obtained. This method can eliminate the initial attitude deviation easily in addition to disturbance rejection. The results show the good performance and superiority of the proposed method compared to some other thrusting maneuver methods.

The main objective of this research is the development of an isolation and detection interacting multiple model (IMM) algorithm for partial (soft) or total (hard) reaction wheel failures in the spacecraft attitude control system (ACS). Based on different scenarios and assumptions we develop healthy models of the ACS under various operating conditions and construct faulty modes with respect to temperature changes, power supply line voltage changes, torque variations, and unexpected current variations in each motor axis windings of the reaction wheels. Once a fault mode is detected and isolated we will be able to develop in certain conditions the recovery procedure through feedback control strategies for the ACS.

Closed-loop attitude steering is a concept for implementing an attitude trajectory by using a conventional quaternion error feedback controller to track the time-varying attitude reference, rather than to simply regulate to a desired orientation. This is done by sampling the reference input and executing the maneuver as a sequence of closely spaced regulating commands that are read out from the spacecraft’s command buffer. The idea has been employed in practice to perform zero-propellant maneuvers on the International Space Station and minimum-time maneuvers on NASA’s TRACE space telescope as well as NASA’s Lunar Reconnaissance Orbiter (LRO). A challenge for operational implementation of the idea is the limited capacity of a space vehicle’s command storage buffer, which is normally not designed with attitude tracking in mind. One approach to mitigate the problem is to downsample-and-hold the attitude commands so that the attitude control system (ACS) regulates to a series of waypoints. This article explores the waypoint following dynamics of a quaternion error feedback control law for such an approach. It is shown that downsample-and-hold induces a ripple between downsamples that causes the satellite angular rate to significantly overshoot the desired limit. Analysis in the z-domain is carried out in order to understand the phenomenon. An interpolating Chebyshev-type filter is proposed that allows the desired attitude trajectory to alternatively be encoded in terms of a small set of filter coefficients. Using the interpolating filter, the continuous-time reference trajectory can be reconstructed and issued at the ACS rate but with significantly reduced memory requirements. The ACS of the LRO is used as an example to illustrate the behavior of a practical ACS.

The Kistler K-l is designed to be a fullyreusable, two stage launch vehicle for economical delivery of small satellite ,payloads to low earth orbit. Greatest efficiency and, hence, lowest cost are achieved by flyback of both vehicle stages to the near vicinity of the launch site. The second stage Orbital Vehicle (OV), after deploying the payload and performing necessary phasing maneuvers, performs a deorbit burn to achieve the desired trajectory conditions at atmospheric entry. After entering the atmosphere, the OV is steered aerodynamically until reaching the deployment point . for a stabilization parachute. Subsequently, drogue and main parachutes complete deceleration of the vehicle for landing orrairbags. This paper presents the design of an atmospheric guidance algorithm for bank-to-turn steering of the OV prior to deployment of the stabilization parachute. Reentry guidance targets a desired geographic position for deployment of the drogue parachute. The algorithm employs a numerical predictor/corrector technique to compute the bank angle and the start time of a single bank reversal required to null the predicted target position miss. Aerodynamic loads and heating are limited implicitly by selection of deorbit target conditions for reentry trajectory shaping. Results obtained using the Draper Laboratory K-l Integrated Vehicle Simulation illustrate guidance performance under nominal and dispersed conditions. * Senior Member of the Technical Staff, Member AIAA. Introduction The K-l Orbital Vehicle (OV) is cylindrical in shape, with a blunt nose and a flared skirt on the aft end (Figure 1). The vehicle trims at an angle of attack of approximately 11 to 12 degrees over the Mach range experienced during reentry. This results in a lift-todrag ratio ranging from 0.1 to 0.15. The only control effecters available are 12 small attitude control system (ACS) jets that are used to provide three-axis attitude stabilization and control. Due .to jet size limitations, alteration of the vehicle angle of attack from the trim condition for purposes of trajectory control is not feasible. Thus, active control of the vehicle trajectory is possible via bank maneuvers only. Given the vehicle mass characteristics and ACS jet properties, the bank acceleration is limited to approximately 2.5 deg/sec/sec. A maximum sustainable bank rate of approximately 15 deg/sec is possible; however, 10 deg/sec is used as a practical limit to avoid excessive Euler coupling (and, hence, fuel usage). Figure 1. Kistler K-l Orbital Vehicle (OV) Copyright 01999 by The Charles Stark Draper Laboratory, Inc. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission: 1275 (c)l999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization. When the Mach number decreases to 2.5, a stabilization parachute is --deployed to stabilize the vehicle about the pitch and yaw axes in the low supersonic and transonic regimes. When a specified altitude is reached a drogue parachute and main parachutes are deployed to decelerate the vehicle for subsequent landing. on airbags. The landing site may be located either at the Nevada Test Site (NTS) or in Woomera, Australia. The key requirement levied on the reentry guidance is to steer the vehicle to the proper conditions for parachute deployment to achieve the landing target position with a given accuracy. This must be accomplished without exceeding specified aerodynamic loading and heating rate limits (8 g longitudinal, 4 g lateral, and 58 BTU/sq ft/sec, respectively). ACS fuel usage must be limited in the process. These requirements must be satisfied in the presence of potentially significant known and unknown dispersions in vehicle mass and aerodynamics properties, deorbit trajectory, and atmospheric conditions. An adaptive atmospheric reentry guidance algorithm has been designed which, as part of an overall guidance, navigation, and control subsystem, addresses the above vehicle characteristics and mission design requirements. The algorithm uses a numerical predictor/corrector technique to solve for a bank angle magnitude and the start time of a single bank reversal that will null the predicted target position miss. Numerical predictor/corrector guidance has been suggested previously for application to aerobraking and aerocapture.“2’3 -While computationally intensive, this technique allows the use of relatively high-fidelity models of vehicle and environmental characteristics, with ,correspondingly increased accuracy over analytic techniques. This paper presents a discussion of the decision’ process leading to the guidance design in addition to a detailed description of the key elements of ‘. the guidance algorithm. Details of the flight software implementation of the guidance algorithms are included where beneficial. Finally, simulation results are supplied to illustrate guidance performance. under nominal and dispersed-conditions. Reentry Dynamics Figure 2 is a free-body diagram of the OV during atmospheric reentry. The lift and drag accelerations experienced by the vehicle are expressed as where. . p f atmospheric density Vd = velocity relative to air mass C, , C, = vehicle drag and lift coefficients

H∞ control option for a combined energy and attitude control system

This paper presents tracking strategies for the attitude dynamics of a rigid body that are global on the configuration space SO(3) and semiglobal over the phase space SO(3)×ℝ3. It is well known that global attractivity is prohibited for continuous attitude control systems on the special orthogonal group. Such topological restriction has been dealt with either by constructing smooth attitude control systems that exclude a set of zero measure in the region of attraction or by introducing discontinuities in the control input. This paper proposes nonmemoryless attitude control systems that are continuous in time, where the region of attraction guaranteeing exponential convergence completely covers the special orthogonal group. This provides a new framework to address the topological restriction in attitude controls. The efficacy of the proposed methods is illustrated by numerical simulations and an experiment.

The role of small inexpensive satellites continuously grows in the modern space exploration. Their use can significantly reduce the cost of the mission. However, the control of such satellites is a challenge. One of the major issues here is that such satellites usually do not possess a complex attitude control system, and three-axis stabilization might be unavailable. As a consequence, the thrust vector of the orbit control system cannot be arbitrary oriented in space and rather involved mathematical methods are needed in order to compensate the control system’s simplicity.

Miniaturized attitude control system for nanosatellites