ALGORITHMS AND OPTIMAL CONTROL FOR SPACECRAFT MAGNETIC ATTITUDE MANEUVERS

This study focused on providing applicable control solutions for spacecraft magnetic attitude control system. Basically, two main lines are pursued; first, developing detumbling control laws and second, an improvement in the three-axis attitude control schemes by extending magnetic rods activation time. Spacecraft, after separation from the launching mechanism, experiences a tumbling phase due to an undesired angular momentum. In this study, we present a new efficient variant of the B-dot detumbling law by introducing a substitute of the spacecraft angular velocity, based on the ambient magnetic field data. This B-dot law preserves the orthogonality, among the applied torque, dipole moment and magnetic field vectors. Most of the existing variants of the B-dot law in the literature don\u27t preserve this orthogonality. Furthermore, the problem of minimum-time spacecraft magnetic detumbling is revisited within the context of optimal control theory. Two formulations are presented; the first one assumes the availability of the angular velocity measurements for feedback. The second formulation assumes the availability of only the ambient magnetic field measurements in the feedback; the latter is considered another optimal-based B-dot law. A reduction in detumbling time is fulfilled by the proposed laws along with less power consumption for the proposed B-dot laws. In magnetic attitude maneuvers, magnetic rods and magnetometers usually operate alternatively, to avoid the magnetic rods\u27 noise effect on magnetometers measurements. Because of that, there will be no control authority over the spacecraft during the magnetometer measurement period. Hence longer maneuver times are usually experienced. In this study, a control scheme that enables the extension of the magnetic rods’ activation time is developed, regardless of the attitude control law. The key concept is replacing the real magnetic field measurement by a pseudo measurement, which is computed based on other sensors measurements. By applying a known command to the spacecraft and measuring the spacecraft response, it is possible to compute the ambient magnetic field around the spacecraft. The system mathematical singularity is solved using the Tikhonov regularization approach. Another developed approach estimates the magnetic field, using a relatively simple and fast dynamic model inside a Multiplicative Extended Kalman Filter. A less maneuver time with less power consumption are fulfilled. These control approaches are further validated using real telemetry data from CASSIOPE mission. This dissertation develops a stability analysis for the spacecraft magnetic attitude control, taking into consideration the alternate operation between the magnetic rods and the magnetometers. It is shown that the system stability degrades because of this alternate operation, supporting the proposed approach of extending the operation time of the magnetic rods

Smart attitude control system for small satellites

The attitude control system is one of the most important systems for satellites, which is essential for the satellite's detumbling, pointing, and orbital maneuver. The conventional attitude control system consists of magnetorquers, reaction wheels, and thrusters. Among these actuators, magnetorquers are widely used for satellite detumbling and attitude control, especially for small satellites and CubeSats. It consumes zero propellant compared with thrusters and has a high chance of survival compared with the reaction wheel as it does not contain any moving parts, which makes them last longer in harsh environments. Conventional magnetorquers utilize air or soft magnetic materials, e.g., iron and alloys, as core, and the magnetic field is generated by feeding the electric current to the wrapped solenoid. Due to the power limit of the small satellites, the magnetic field strength is strictly limited, and the continuous current supply results in massive energy consumption for detumbling and other attitude adjustment missions. The long copper wire of the solenoid will also result in high resistance and generate significant heat. To improve the current design and overcome the proposed drawbacks, a novel electro-permanent magnetorquer has been designed and developed in this thesis as one actuator of the attitude control system. Unlike conventional magnetorquers, the electro-permanent magnetorquer utilizes hard magnetic materials as the core, which can maintain the magnetization when the external magnetic field is removed, to generate the required magnetic field. A special driving circuit is designed to generate the desired dipole moment for the magnetorquer, and the components used for the circuit are carefully selected. The experiments show that the electro-permanent magnetorquer can generate 1.287 Am2 dipole moment in either direction. The magnetorquer works in pulse mode to adjust the dipole moment, and it requires around 0.75 J energy maximum per pulse. A single-axis detumbling experiment has been conducted using only one torque rod on the air-bearing table inside an in-house manufactured Helmholtz cage. The experiment results show that the magnetorquer can detumble the air bearing table with 0.0612 kgm2 moment of inertia from an initial speed of around 27°/s to zero within 800s, and total energy of 82.92 J was consumed for the detumbling experiment. A single torque rod single-axis pointing experiment has been conducted with a sliding mode controller on the same platform. The results show that a single torque rod can achieve the target angle and maintain the error discrepancy within the ±0.4° boundary under a specific system configuration. A micro air-fed magnetoplasmadynamic thruster has been designed and tested as another attitude control system actuator. The thruster is a miniaturized electric propulsion system based on the conventional full-scale magnetoplasmadynamic thruster that operates at hundreds of kilowatts. The thruster is designed and tested using normal air as the propellant under the pulse operation mode on a calibrated micro-force measurement thruster stand. The experiments revealed that the thruster could generate a 34.534 µNs impulse bit with an average power input of 1.857 ± 0.0679 W and thrust to power ratio of 8.266 µN/W. The specific impulse is calculated to be 2319 s with a thruster efficiency of 9.402%, which is quite competitive compared with other solid-state and liquid-fed pulse-mode thrusters. This paper presents the design and test results for the thruster under a low power level, as well as an analysis of its problems and limitations with corresponding future research and optimization directions noted at the end. The electro-permanent magnetorquer as a payload of the CUAVA-2 satellite mission has been introduced in this thesis. The design considerations and adjustment based on the requirement of the CUAVA-2 has been introduced in detail. A simple sliding mode controller has been developed to achieve three-axis attitude control using both electro-permanent magnetorquer and the micro air-fed magnetoplasmadynamic thruster. The controller's performance has been tested using MATLAB-based simulation with the experimentally obtained performance parameters and some assumptions. The results show that the smart attitude control system can achieve ±0.005° pointing error discrepancy with the help of both actuators

Attitude Determination and Control Subsystem Design for a CubeSat

This project continues the design and testing of the Attitude Determination and Control Subsystem (ADCS) for a nano-satellite. The primary mission objective is solar X-ray spectroscopy using the Sphinx-NG instrument, which requires that the CubeSat fly in a high-altitude, polar, sun- synchronous orbit pointing to the sun with 1-2 degrees of accuracy. The ADCS requires gyroscopes, sun sensors, and a magnetometer for attitude determination. Attitude control is executed using magnetorquers as actuators. This project focused on the analysis of attitude determination algorithms and control policies to select the most efficient and accurate methods. After method selection, simulations of the ADCS were conducted, and research was performed concerning hardware testing for the ADCS

운영 시나리오를 고려한 저궤도 큐브위성의 자세결정 및 제어에 관한 연구

학위논문 (석사)-- 서울대학교 대학원 : 기계항공공학부, 2016. 8. 기창돈.SNUGLITE (Seoul National University GNSS Laboratory satellITE), a cubesat scheduling for launch in 2017, must maintain nadir pointing attitude control in order to obtain mission data successfully and send them to ground station using S-band antenna. This thesis will examine the ADCS (Attitude Determination and Control System) algorithm and how the proposed algorithm satisfies the ADCS requirements for this particular operation scenario. The ADCS is composed of 3-axis MEMs gyroscope sensor, 2-axis coarse photodiode type sun sensors, 3-axis MEMs magnetometer, dual frequency GPS receivers, and 3-axis magnetic torquers. In the simulation environment, studies were done on the eclipse and various disturbance torques, including gravity gradient torque, solar radiation torque, and aerodynamic torque. The LQG (Linear Quadratic Gaussian) controller has been chosen for ADCS algorithm for SNUGLITE. Furthermore, this thesis will provide SILS (Simulation In the Loop System)—which takes low earth orbit environments into account—to verify the proposed ADCS algorithm. Due to the fact that the precise launch schedule of the cubesat has not yet been decided, the orbit elements have been chosen for the worst case scenario: it will provide a maximum eclipse time in altitude of 600[km] from the circular suns synchronous orbit. In order to monitor the low earth orbit environment which has a direct effect on the attitude dynamics, we developed a low earth orbit simulator that is comprised of attitude dynamics and orbit dynamics which can influence each other. We used two computers, one for the ADCS algorithm and one for the orbit environment, and transmitted data using serial communication. Thus, the team of scholars after us could use this SILS we developed to further the research on PILS (Processor In the Loop System). The operation scenario consists of two parts. The first step starts from deployment of cubesat from P-POD (Poly Picosatellite Orbital Deployer), and detumbling using B-dot control. Second, the use of LQG controller for nadir pointing control. However, the second part of the operation scenario is also divided into two segments. In the first phase, only GPS is used as a payload but magnetic boom is not deployed, while in the second phase, magnetic boom is used for the earthquake mission. After evaluating the simulation results, we have come to a conclusion that all the ADCS requirements were met. For instance, the attitude estimation errors were less than 5 [deg] in eclipse and less than 2 [deg] per day. In addition, the attitude control errors were less than 10 [deg] in eclipse and less than 5[deg] per day. Finally, the ADCS algorithm enabled the cubesat to turn over even in an up-side-down position. In summary, this thesis developed and verified the ADCS algorithm which was based on the Matlab by using the LQG controllermoreover, it offers the space environment simulator which could be used for the PILS (Processor In the Loop System) study in the future. We expect these results will contribute to making SNUGLTIEs mission a success.I. Introduction 1 1. Motivation and Objectives 1 Mission for SNUGLITE 1 Necessity for ADCS (Attitude Determination and Control System) 2 Necessity for SILS (Software In the Loop System) 3 2. Previous Work 6 3. Contents and Method 10 4. Contribution 11 II. Space Environment 13 1. Modeling of Orbit 14 Sun synchronous orbit 14 Definition of coordinate framen 15 Disturbance torques 17 Earth Magnetic/Sun Model 21 2. Modeling of Cubesat 24 SNUGLITE specification 25 Attitude dynamics 26 III. ADCS Algorithm 29 1.ADCS requirements 29 2. Operation Scenario 31 3. Attitude Estimation 33 TRIAD 33 EKF 34 4. Attitude Control 43 B-dot 44 LQG 45 IV. Simulation 50 1. Simulation Configuration 50 2. Simulation Environment 52 3. Simulation Result 53 V. Conclusion 64 VI. References 65 요약 68Maste

Development of a ground testing facility and attitude control for magnetically actuated nanosatellites

Growing popularity of the highly capable small- and nano-satellites, driven by components miniaturization, face new technological challenges and at the same time provides new opportunities for the whole space sector. Low cost of nanosatellites launches make them accessible. Reliability is an exigency: especially challenging is design and testing of Attitude and Determination Control Systems (ADCS). Demand for nanosatellitesdedicated attitude control algorithms and careful performance assessment of the spacecrafts motivates the research work presented in this thesis. In the first part of the manuscript, development and assessment of the three degreesoffreedom ADCS testbed for nanosatellites testing is described. The facility was developed within the Microsatellites and Space Microsystems Lab at University of Bologna, and designed to meet strict low-cost requirements. The facility includes several integrated subsystems to simulate the on-orbit environment: i) an air-bearing based, three degree of freedom platform with automatic balancing system, ii) a Helmholtz , iii) a Sun simulator, and iv) a metrology vision system . Experimental assessment of the subsystems guarantee necessary level of performance. Control law design for smallsats is addressed in the second part. Limited power availability and reliability makes magnetic actuation particularly suited for ADCS design, but, the control system faces inherent underactuation. To overcome the intrinsic limits of existing control designs, a novel approach to the three-axis attitude control of a magnetically actuated spacecrafts is proposed, based on hybrid systems theory. A local H-inf regulator with guaranteed performance and a global nonlinear controller used for ensuring global stability and robustness, are combined. Hybrid control theory is employed to develop a mixed continuous-discrete controller able to switch between different feedbacks. Analytical results are verified by means of realistic numerical simulations: errors on the state comply with the computed bounds and stability is guaranteed

Analysis and design of the 3Cat-8 pointing control system for high-volume data downlink

The objective of this bachelors thesis is to conduct a comprehensive analysis of the communication subsystems in the UHF, S, and X bands of the 3Cat-8 mission. This mission aims to develop a satellite under the CubeSat standard that allows the study of different physical phenomena, such as the effects of auroral emissions on wave propagation through the ionosphere. Furthermore, this project also includes the design of an S-band patch antenna, both to understand the design of a space application antenna and to implement it on the same satellite and compare it with other commercial options. In addition, the polar orbit described by the satellite results in limited opportunities for data downlink, making it necessary to design and dimension magnetic actuators to enable the satellite to point towards the ground station. This document addresses the study, design, and manufacturing of these elements in order to enable the satellite to aim its antennas at the ground station. Overall, this work represents a significant contribution to the 3Cat-8 mission. Throughout this project, various activities encompassing the fields of aerospace engineering and telecommunications have been carried out

Flight Mechanics/Estimation Theory Symposium, 1992

This conference publication includes 40 papers and abstracts presented at the Flight Mechanics/Estimation Theory Symposium on May 5-7, 1992. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers

Flight Mechanics/Estimation Theory Symposium, 1994

This conference publication includes 41 papers and abstracts presented at the Flight Mechanics/Estimation Theory Symposium on May 17-19, 1994. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers

Development and characterization of a standardized docking system for small spacecraft

Since the first mating manoeuvre in space, performed in 1966, many different docking mechanisms were developed, mainly for large manned spacecraft. The few systems recently conceived for small satellites have never been verified in space nor scaled to CubeSat size. In the near future, small spacecraft docking procedures could acquire great importance due to the need to share resources between clusters of low-weight and low-cost vehicles: in fact, small spacecraft market is rapidly growing, focusing on commercial low risk application, low budget scientific and educational missions. In this context, this document presents a novel docking mechanism to provide small spacecraft with the ability to join and separate in space, to realize multi-body platforms able to rearrange, be repaired or updated, thus overcoming the actual on board limitations of single small-scale satellites. As for now, the few proposed docking ports present (1) simple probe-drogue interfaces, unable to dock with same-gender ports, or (2) androgynous geometries, that can overcome that problem, but usually employing complex and non-axis-symmetric latches to perform the docking manoeuvre, that would demand robust and stringent navigation and control systems. The proposed solution overcomes the aforementioned drawbacks, using a semi-androgynous shape-shifting mechanism that actuating one interface changes the port into a “drogue" configuration, letting the other port penetrate it and closing around to create a solid joint. The mechanism design through the requirement definition and a trade-off between different concepts is presented, followed by the analysis of the dynamic behaviour of the selected solution, with particular attention to two aspects, i.e. the loads transmitted between the mating ports and the alignment tolerances requested to perform successful docking manoeuvres. Such analysis led to the definition of an instrumented prototype to verify the solution through simple validation tests, which demonstrated the mechanism operations and defined the alignment ranges, that lie in the range of +- 15 mm and up to 6 degrees. Last, a comparison with SPHERES UDP is presented, as part of the activities performed during a visit period at MIT Space Systems Laboratory