Development of attitude determination control system for nanosatellites

This thesis presents a design of the system for the first stage of a mission in the NanoSat Lab, the 3Cat-8. There is an special focus on the control algorithm and the sizing of the actuators

A Large Scale Simulation of Satellites Tracking Vessels and Other Targets

This research outlines the design of a large scale simulation of satellites tracking large amounts of dynamic targets. The use of such a simulation is presented and current solutions available are presented. The research sets out a list of objectives to meet by creating an application programming interface (API) that have the requirements of being efficient, scalable, flexible, and easy to use for the implementer. Methods of creating sections of the simulation such as the attitude motion of a satellite based on the physical characteristics of nanosatellites is explored and developed. The creation of targets that are contained only on certain land features are also developed and tested. The objectives set out are tested by creating a simulation using the API developed and the results are presented

Hybrid Fuzzy Logic and Extremum Seeking Attitude Control of Solar Sail Spacecraft

This thesis explores four controllers applied to the attitude control of a solar sail craft with control masses with the goal of showing benefits over more standard control schemes. The controllers examined in this paper are: 1) a PID controller that incorporates a discrete extremum seeking algorithm, 2) a type-1 fuzzy logic controller that incorporates a discrete extremum seeking algorithm, 3) a type-1 fuzzy logic controller, and 4) a type-2 fuzzy logic controller. The first two controllers are examined for their ability to quickly converge to a set of optimal gains over time. The latter two controllers are evaluated for their ability to maintain stability with respect to model uncertainty and sensor noise. The four controllers discussed in this paper are compared against other control techniques that have already been shown in previously published literature to provide good control performance when applied to this system

Robust algorithms for the identification and control of Android-powered quadcopters

The focus of this thesis is on modeling and control of a non-real time, Android operated quadcopter of type dji F450 Flame Wheel without having a concrete knowledge about the system's dynamics and parameters. The quadcopter is equipped with an onboard non-rooted Android smartphone, which serves as both the controller and the IMU in the system. The reference command signals are generated by another user-held Android smartphone which defines the desired orientation of the quadcopter. Due to the fact that default Android implementation is not real-time, the measurements of both Android phones are subject to significant latencies resulting in asynchronous data. To obtain a model of the system, a comprehensive system identification study of the quadcopter's rotation dynamics using grey box model and Euler's equations of rigid body is introduced in the thesis. It also introduces two novel algorithms for obtaining an initial guess for the inertia matrix using convex optimization despite the presence of large number of local minimizers in the original prediction error problem. It shows how sensitive the process is to the initial guess of the model's parameters. A detailed comparison of the relevant estimates is also shown. The control laws were implemented on the onboard Android device, which reads the asynchronous built-in sensors measurements and generates the control signals required to steer the quadcopter and obtain the desired orientation defined by the user-held device. Two control laws were developed, an advanced model-free PID controller that accounts for the non-uniformly distributed data, the windup effect, and the derivative kick, and a model-based LQI controller. Both control approaches were able to stabilize the quadcopter despite the data asynchronousity and model uncertainty, and were validated and tested empirically and through simulation. The thesis also introduces a novel approach of optimizing the PID controllers gains based on the jacobian matrix. The optimization problem tends to be poorly conditioned for such systems. Hence, the novel scaling technique improves the conditioning of the optimization problem and obtains better minimizers. The efficiency of the proposed algorithm is evaluated through simulation. Furthermore, a detailed study on the effect of the cost function selection and model uncertainty on the optimization process is shown

Attitude regulation with bounded control in the presence of large disturbances with bounded moving average

The attitude regulation problem with bounded control for a class of satellites in the presence of large disturbances, with bounded moving average, is solved using a Lyapunov-like design. The analysis and design approaches are introduced in the case in which the underlying system is an integrator and are then applied to the satellite attitude regulation problem. The performance of the resulting closed-loop systems are studied in detail and it is shown that trajectories are ultimately bounded despite the effect of the persistent disturbance. Simulation results on a model of a small satellite subject to large, but bounded in moving average, disturbances are presented

A scalable, portable, FPGA-based implementation of the Unscented Kalman Filter

Sustained technological progress has come to a point where robotic/autonomous systems may well soon become ubiquitous. In order for these systems to actually be useful, an increase in autonomous capability is necessary for aerospace, as well as other, applications. Greater aerospace autonomous capability means there is a need for high performance state estimation. However, the desire to reduce costs through simplified development processes and compact form factors can limit performance. A hardware-based approach, such as using a Field Programmable Gate Array (FPGA), is common when high performance is required, but hardware approaches tend to have a more complicated development process when compared to traditional software approaches; greater development complexity, in turn, results in higher costs. Leveraging the advantages of both hardware-based and software-based approaches, a hardware/software (HW/SW) codesign of the Unscented Kalman Filter (UKF), based on an FPGA, is presented. The UKF is split into an application-specific part, implemented in software to retain portability, and a non-application-specific part, implemented in hardware as a parameterisable IP core to increase performance. The codesign is split into three versions (Serial, Parallel and Pipeline) to provide flexibility when choosing the balance between resources and performance, allowing system designers to simplify the development process. Simulation results demonstrating two possible implementations of the design, a nanosatellite application and a Simultaneous Localisation and Mapping (SLAM) application, are presented. These results validate the performance of the HW/SW UKF and demonstrate its portability, particularly in small aerospace systems. Implementation (synthesis, timing, power) details for a variety of situations are presented and analysed to demonstrate how the HW/SW codesign can be scaled for any application

Guidance, Navigation, and Control of Small Satellite Attitude Using Micro-Thrusters

M.S. University of Hawaii at Manoa 2016.Includes bibliographical references.In this study, a new and automated Navigation, Guidance and Control system is designed, analyzed, simulated and tested for small satellites. As is known, this system represents the primary unit of on-board control of a flight vehicle. It consists of a set of system software algorithms and hardware elements, including various sets of sensors and electronics depending on the type of the vehicle. This study is focused on small satellites, which are becoming one of the primary tools for a wide range of low Earth and deep space missions. The Navigation subsystem has been described in terms of its sensors and filtering technique, known as the Extended Kalman Filter. This subsystem provides the estimates of the satellite’s state vector. It is assumed that this vehicle’s Navigation subsystem includes GPS receiver, and accelerometer and gyro, which are considered as Inertial measurement Unit (IMU) component subsystems. The Guidance subsystem provides guidance commands for satellite’s actuators, which are assumed to include a set of micro-thrusters. The Control subsystem provides control commands for increments of torque of actuation. This study deals with the development, design and integration of the Navigation, Guidance and Control (known as GNC) subsystems into a unique framework that can be executed on-board in real time to perform satellite attitude maneuvers. The main focus is on the development of Guidance subsystem functions and algorithms. These functions, in particular, include attitude angles, angular rates and coefficients. The Guidance subsystem provides commanded angular acceleration based on a fourth-order polynomial with respect to time, which was used for lunar-descent trajectory guidance during the Moon landing maneuvers of Apollo Landers. The difference in the utility of this polynomial law in Apollo missions and this work is that in those missions this polynomial was used for trajectory guidance using numerically integrated trajectories as reference solutions. In this work, this polynomial is used to compute attitude guidance commands using a simple PD controller as an analytic reference attitude profile. The novelty of this work is that this polynomial law is formulated and implemented for the first time for real-time and on-board attitude guidance and control using a set of microthrusters as part of the integrated GNC system. Another element of novelty is associated with targeting. A real-time targeting procedure implies on-board computations of the target states and the time remaining to achieve the target state from the current state. In this work, the target state includes Euler angles and their rates. As such, the targeting is considered as an integral and critical part of the guidance function. The guidance command is computed only after computations of the target state and is the explicit function of this state. Therefore, the proposed guidance function is considered as the ob-board target-relative attitude guidance. The performance of the proposed GNC system has been demonstrated by two illustrative examples. In the first example, the satellite is guided to orient itself to its target position. In the second example, the satellite is guided to perform two consecutive rotational maneuvers, detumbling and reorientation, to achieve a desired attitude. The numerical simulation parameters and its results are illustrated by various plots and qualitative analysis of the relationships between the satellite’s state and guidance parameters. The list of references and appendix with necessary formulas and figures are provided

Design and verification of Guidance, Navigation and Control systems for space applications

In the last decades, systems have strongly increased their complexity in terms of number of functions that can be performed and quantity of relationships between functions and hardware as well as interactions of elements and disciplines concurring to the definition of the system. The growing complexity remarks the importance of defining methods and tools that improve the design, verification and validation of the system process: effectiveness and costs reduction without loss of confidence in the final product are the objectives that have to be pursued. Within the System Engineering context, the modern Model and Simulation based approach seems to be a promising strategy to meet the goals, because it reduces the wasted resources with respect to the traditional methods, saving money and tedious works. Model Based System Engineering (MBSE) starts from the idea that it is possible at any moment to verify, through simulation sessions and according to the phase of the life cycle, the feasibility, the capabilities and the performances of the system. Simulation is used during the engineering process and can be classified from fully numerical (i.e. all the equipment and conditions are reproduced as virtual model) to fully integrated hardware simulation (where the system is represented by real hardware and software modules in their operational environment). Within this range of simulations, a few important stages can be defined: algorithm in the loop (AIL), software in the loop (SIL), controller in the loop (CIL), hardware in the loop (HIL), and hybrid configurations among those. The research activity, in which this thesis is inserted, aims at defining and validating an iterative methodology (based on Model and Simulation approach) in support of engineering teams and devoted to improve the effectiveness of the design and verification of a space system with particular interest in Guidance Navigation and Control (GNC) subsystem. The choice of focusing on GNC derives from the common interest and background of the groups involved in this research program (ASSET at Politecnico di Torino and AvioSpace, an EADS company). Moreover, GNC system is sufficiently complex (demanding both specialist knowledge and system engineer skills) and vital for whatever spacecraft and, last but not least the verification of its behavior is difficult on ground because strong limitations on dynamics and environment reproduction arise. Considering that the verification should be performed along the entire product life cycle, a tool and a facility, a simulator, independent from the complexity level of the test and the stage of the project, is needed. This thesis deals with the design of the simulator, called StarSim, which is the real heart of the proposed methodology. It has been entirely designed and developed from the requirements definition to the software implementation and hardware construction, up to the assembly, integration and verification of the first simulator release. In addition, the development of this technology met the modern standards on software development and project management. StarSim is a unique and self-contained platform: this feature allows to mitigate the risk of incompatibility, misunderstandings and loss of information that may arise using different software, simulation tools and facilities along the various phases. Modularity, flexibility, speed, connectivity, real time operation, fidelity with real world, ease of data management, effectiveness and congruence of the outputs with respect to the inputs are the sought-after features in the StarSim design. For every iteration of the methodology, StarSim guarantees the possibility to verify the behavior of the system under test thanks to the permanent availability of virtual models, that substitute all those elements not yet available and all the non-reproducible dynamics and environmental conditions. StarSim provides a furnished and user friendly database of models and interfaces that cover different levels of detail and fidelity, and supports the updating of the database allowing the user to create custom models (following few, simple rules). Progressively, pieces of the on board software and hardware can be introduced without stopping the process of design and verification, avoiding delays and loss of resources. StarSim has been used for the first time with the CubeSats belonging to the e-st@r program. It is an educational project carried out by students and researchers of the “CubeSat Team Polito” in which StarSim has been mainly used for the payload development, an Active Attitude Determination and Control System, but StarSim’s capabilities have also been updated to evaluate functionalities, operations and performances of the entire satellite. AIL, SIL, CIL, HIL simulations have been performed along all the phases of the project, successfully verifying a great number of functional and operational requirements. In particular, attitude determination algorithms, control laws, modes of operation have been selected and verified; software has been developed step by step and the bugs-free executable files have been loaded on the micro-controller. All the interfaces and protocols as well as data and commands handling have been verified. Actuators, logic and electrical circuits have been designed, built and tested and sensors calibration has been performed. Problems such as real time and synchronization have been solved and a complete hardware in the loop simulation test campaign both for A-ADCS standalone and for the entire satellite has been performed, verifying the satisfaction of a great number of CubeSat functional and operational requirements. The case study represents the first validation of the methodology with the first release of StarSim. It has been proven that the methodology is effective in demonstrating that improving the design and verification activities is a key point to increase the confidence level in the success of a space mission

헬름홀츠 케이지를 이용한 저궤도 큐브위성 자세결정 및 제어시스템의 HILS 검증

학위논문 (석사)-- 서울대학교 대학원 : 공과대학 기계항공공학부, 2019. 2. 기창돈.본 논문에서는 자기토커를 구동기로 탑재한 저궤도 큐브위성에 대해, 지구지향 자세유지를 위한 자세결정 및 제어시스템의 성능검증 기법을 제안한다. 반작용휠을 탑재한 일반적인 위성과 달리, 공간적 제약을 가지는 큐브위성의 효과적인 운용을 위해 단순하고 저중량, 저전력의 자기토커만을 구동기로 탑재하여 지구지향 자세유지를 수행한다. 이를 위해, 먼저 큐브위성 자세에 대한 운동방정식을 중력구배토크, 그리고 자기토커의 쌍극자모멘트와 지구 자기장의 외적으로 산출되는 입력 토크에 대해 표현하고, 시스템에 유입되는 불확실성을 잡음원으로 취급하여 선형 시스템 모델을 얻는다. 또한, 태양과 자기장 모델로부터 기준벡터를 정의하고 태양 센서와 자기장 센서, 그리고 자이로스코프 측정치를 융합하여 큐브위성의 자세를 추정하는 확장 칼만 필터를 구성한다. 여기에, 주어진 선형 시스템과 입력에 대해 비용함수를 구성하고 최적해를 산출함으로서 지구지향 자세를 유지하는 LQR 제어기를 설계한다. 제안된 시스템의 성능을 검증하기 위해, 지상환경의 다양한 제약조건을 고려한 HILS가 수행되어야 한다. 하지만, 자기토커만 탑재한 큐브위성은 작은 입력토크, 지구 자기장 크기에 비례한 출력결정, 비연성 제어구동 등의 구동성능 한계로 인해 지상환경에서 전술한 시스템의 효과적인 성능 검증이 어렵다. 지금까지 큐브위성 자세결정 및 제어시스템의 성능검증은 대부분 운용의 안정성 확보를 위한 각속도 안정화, 지구 자기장 정렬에 의존한 수동제어 방식, 그리고 반작용휠을 활용한 지구지향 자세유지의 HILS가 연구되어 왔다. 자기토커만을 탑재한 자세결정 및 제어시스템의 HILS가 제안된 바가 있으나, 지상환경의 제약조건으로 인하여 성능검증에 한계가 있었다. 이와 달리, 제안되는 HILS는 기존방법의 한계를 보완하여 다양한 오차가 내포하는 환경에서 자기토커만을 구동기로 탑재한 큐브위성 자세결정 및 제어시스템의 성능검증에 목표를 두고 있다. 따라서 본 논문에서는 자기토커의 외부자기장 크기에 비례한 출력특성에 착안하여, 헬름홀츠 케이지를 이용한 큐브위성 자세결정 및 제어시스템의 HILS 검증기법이 제안된다. 즉, 지상환경에서 통계적 오차특성을 내포하는 자기장으로 인한 추정 성능저하와 외란에 취약한 자기토커의 작은 입력토크로 인한 제어 성능저하 문제를 헬름홀츠 케이지로부터 생성된 자기장 벡터를 제어함으로써 해결하고자 한다. 이를 위해, 비오-샤바르 법칙을 활용하여 헬름홀츠 케이지의 전류-자기장 관계를 모델링하고 큐브위성을 포함한 공간의 자기장 균일성을 확보할 수 있는 헬름홀츠 케이지를 제작한다. 또한 제시된 헬름홀츠 케이지의 전류-자기장 모델로부터 전달함수를 근사하고 고전제어기법을 활용하면 헬름홀츠 케이지의 자기장 벡터 제어기를 손쉽게 설계할 수 있다. 여기에, 헬름홀츠 케이지 내부공간에 큐브위성을 실에 매달아 단일축 자세결정 및 제어시스템 HILS 환경을 구성하고, 자세결정 및 제어시스템의 HILS 검증을 수행한다. 이때, 실내에 구축된 실험환경에서 GPS 측정치를 산출할 수 없으므로 모사된 태양광과 헬름홀츠 케이지에서 생성되는 자기장의 평균 측정치를 기준벡터로 재정의하여 지구지향 자세를 모사한다. 제시된 HILS 환경에서 헬름홀츠 케이지를 기준으로 좌표계를 정의하면, 지상환경에서 큐브위성의 지구지향 자세유지에 대한 시뮬레이션을 구성할 수 있다. 이러한 시뮬레이션 결과에 근거하여 실제 실험결과와 비교하면, 제안된 자세결정 및 제어시스템의 성능을 해석할 수 있다. 제안된 방법의 유용성을 확인하기 위해, SNUGLITE(Seoul National University GNSS Laboratory satelliTE) 큐브위성의 단일축 자세결정 및 제어시스템 HILS 검증이 제시된다. 제안된 방법은 지상환경에서 큐브위성에 대표적으로 탑재되는 자기토커만을 활용하여, 큐브위성의 지구지향 자세유지를 위한 자세결정 및 제어시스템을 검증할 수 있음을 보인다. 실험결과로부터 기존 방법의 지상환경 제약조건으로 인한 HILS 검증 한계를 헬름홀츠 케이지를 활용하여 보완함을 보인다. 또한, 헬름홀츠 케이지를 활용하지 않는 기존 방법과 비교하여 자세결정 및 제어시스템의 추정 성능 및 제어 신뢰성을 효과적으로 검증할 수 있음을 확인한다. 제안된 HILS 검증기법은 간결성 및 실용성으로 인해 다양한 임무수행을 위한 큐브위성의 자세결정 및 제어시스템 검증에 활용될 것으로 기대된다.In this thesis, Hardware-In-the-Loop Simulation (HILS) verification of attitude determination and control system (ADCS) is addressed for low earth orbit cube-satellite equipped with a magnetorquer only. Unlike ordinary satellites equipped with reaction wheels, only a magnetorquer is mounted on cube-satellite due to spatial constraints. It is a simple, low-weight, and low-power consumption actuator that enables efficient operation of cube-satellite and aims to maintain nadir-pointing control. In order to achieve the objective of the proposed system, firstly, the equations of motion for the cube-satellite is expressed in terms of the gravity gradient torque, and the input torque calculated from the dipole moment of the magnetorquer and geomagnetic field. Then, a linear system model is obtained by interpreting the uncertainty flowing into the system as noise sources. In addition, extended Kalman filter equations are derived to estimate the attitude of the cube-satellite by defining the reference vector from the sun and magnetic field model, and fusing the sun sensor, magnetometer, and gyroscope measurements. Then, LQR controller can be designed to maintain nadir-pointing by calculating the optimal solution from the cost function composed of a given linear system and input. In order to verify the performance of the proposed system, HILS should be performed considering various constraints of the ground environment. However, it is difficult to verify the above-described system in a ground environment due to limitations in the performance of magnetorquer such as small input torque effected by the magnitude of the geomagnetic field, and decoupled input control. So far, it has been studied that the verification of cube-satellite ADCS for the stabilization of angular velocity, the passive control method which depends on geomagnetic field alignment, and HILS using reaction wheels. HILS of ADCS using only magnetorquer also has been proposed, but the performance analysis was limited due to the constraints of the ground environment. In contrast, the proposed HILS is aimed at verifying magnetorquer mounted cube-satellite ADCS that solves the limitations from unknown error factors of the ground environment. Therefore, in this thesis, HILS verification of cube-satellite ADCS using Helmholtz cage is proposed. It is focused on output characteristics proportional to the magnitude of the external magnetic field of the magnetorquer. In other words, it is solved by controlling the magnetic field vector generated from Helmholtz cage that is the degradation of the estimation performance due to the magnetic field including the statistical error characteristic in the indoor environment and the control performance deterioration due to the small input torque of the magnetorquer which is vulnerable to the disturbance. To construct a magnetic field vector from Helmholtz cage, it is designed using the Biot-Savat law to model the current-magnetic field relationship. Also, it is designed to have a size enough to ensure the magnetic field uniformity of the space including the cube-satellite. The magnetic field vector controller of Helmholtz cage can be easily designed by approximating the transfer function from the derived current-magnetic field equation and using the classical control technique. In this case, cube-satellite is suspended in the inner space of Helmholtz cage to perform single axis HILS verification of ADCS. Since the GPS measurement value cannot be calculated in the indoor experiment environment, the nadir-pointing reference vectors are redefined by the mean measurement values of the simulated sunlight and the magnetic field generated from Helmholtz cage. Then, by defining a coordinate system based on the Helmholtz cage in the proposed HILS environment, nadir-pointing control performance can be expected by the computer simulation. Based on these simulation results, the proposed system can be verified by comparison with actual experimental results. To demonstrate the validity of the proposed method, a single axis HILS verification of ADCS using SNUGLITE cube-satellite is presented. The proposed method can verify the performance of nadir-pointing control on the ground by using only the magnetorquer which is typically installed in cube-satellite. It is also confirmed that the estimation performance and the control reliability of ADCS can be verified effectively compared with the method which does not utilize Helmholtz cage. The proposed HILS verification technique is expected to be used for verification of cube-satellite ADCS for various tasks due to its simplicity and practicality.Abstract i Table of Contents v Nomenclature viii List of Figures x List of Tables xiii Chapter 1 Introduction 1 1.1. Brief Review of SNUGLITE Cube-Satellite Project 2 1.1.1. Introduction of SNUGLITE Cube-Satellite 2 1.1.2. SNUGLITE Configuration 4 1.1.3. Former Research of SNUGLITE ADCS 8 1.2. Motivation and Purpose 9 1.3. Literature Survey 11 1.4. Outline of Research 13 1.5. Contribution 15 Chapter 2 Algorithm of Attitude Determination and Control System 17 2.1. Overall System Configuration 18 2.2. Coordinate System 19 2.2.1. Earth-Centered Inertial (ECI) Frame 21 2.2.2. Earth-Centered Earth-Fixed (ECEF) Frame 21 2.2.3. Local Frame 22 2.2.4. Body Frame 23 2.3. Coordinate Transformations 24 2.3.1. Transformation Between ECEF and ECI Frame 25 2.3.2. Transformation Between ECI and Local Frame 25 2.3.3. Transformation Between Local and Body Frame 26 2.4. Dynamics Modeling 28 2.4.1. Nonlinear Equations of Motion 28 2.4.2. Linearized Equations of Motion 32 2.5. Angular Velocity Attenuation for Initial Phase 33 2.6. Attitude Determination Algorithm 34 2.6.1. TRIAD Method 35 2.6.2. Extended Kalman Filter 36 2.7. Attitude Control Algorithm 42 Chapter 3 A Solution for Magnetic Distortion of Cube-Satellite 45 3.1. Magnetometer Calibration for Time-varying Bias 46 3.1.1. Temperature Calibration 46 3.1.2. Current Compensation 50 3.1.3. Hard-Iron and Soft-Iron Compensation 57 3.2. Residual Magnetic Dipole Moment 60 3.3. Simulation 63 3.3.1. Simulation Envirionment 63 3.3.2. Results 69 Chapter 4 Hardware-In-the-Loop Simulation of Attitude Determination and Control System 81 4.1. Helmholtz Cage Design 82 4.1.1. Introduction 82 4.1.2. Mathematical Modeling and Simulation 85 4.1.3. CAD and Assembly 90 4.1.4. Driving Circuit Design 93 4.1.5. Magnetic Field Controller Design 96 4.2. Hardware-In-the-Loop-Simulation 102 4.2.1. HILS Configuration 102 4.2.2. Experiment Results 112 Chapter 5 Conclusion 117 References 120 Abstract in Korean 124Maste