Multidisciplinary Control of a Sparse Interferometric Array Satellite Testbed

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.

Spacecraft attitude control systems rely on reaction wheels as the primary means of precise three-axis attitude control. Faults in these reaction wheels might lead to system instability and, in severe cases, mission failure. This paper presents advanced machine learning-based techniques for the detection, isolation, and identification of reaction wheel faults in spacecraft.The proposed approach leverages advanced data analytics and machine learning algorithms to analyze sensor data from the reaction wheels, enabling early detection of faults and effective isolation of the faulty component and identify the types of faults detected, specifically, voltage, current and temperature faults.Three-axis controlled satellite high-fidelity models are simulated to generate data for both nominal and faulty states of RW. The simulated data is employed with the FDII approach. The generated data is passed into five different machine learning classifiers, the isolation and identification results are verified via cross-validation. The proposed techniques is tested on three defined datasets using the three-orthogonal RW configuration to verify its robustness. The results show that the system has higher isolation and identification accuracy when compared to other studies that used various methodologies.

Design and wide-bandwidth control of large aperture fast steering mirror with integrated-sensing unit

Space environmental effects on materials are very severe and complex because of the synergistic interaction of orbital environments such as high-energy radiation particles, atomic oxygen, micrometeoroids, orbital debris, and ultraviolet irradiation interacting synergistically, along with thermal exposure. In addition, surface degradation associated with contamination can negatively impact optics performance. Materials flight experiments are critical to understanding the engineering performance of materials exposed to specific space environments. Likewise, the spacecraft designer must have an understanding of the specific environment in which a spacecraft will operate, enabling appropriate selection of materials to maximize engineering performance, increase mission lifetimes, and reduce risk. This article will present a methodology for assessing the engineering performance of materials baselined for a specific spacecraft or mission. In addition, an overview of the space environment, from low Earth orbit to interplanetary space, will be provided along with an overview on the effects of the space environment on materials performance. The majority of this article is devoted to materials flight experiments from the European Space Agency (ESA), the Japan Aerospace Exploration Agency (JAXA), and from the National Aeronautics and Space Administration (NASA). Some of the experiments reviewed include ESA's Materials Exposure and Degradation Experiment on the International Space Station (ISS), JAXA's Micro-Particles Capturer and Space Environment Exposure Device experiments on the ISS Service Module and on the ISS Japanese Experiment Module Exposed Facility, and NASA's Long Duration Exposure Facility satellite and the Materials International Space Station Experiment series flown on the exterior of ISS.

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 objective of the research is to investigate an effective method of using servo control through the use of a fast steering mirror and/or acoustic actuators and sensors for optical jitter suppression. The experimental testbed is built in an anechoic chamber at Duke University. Two optical benches are used to isolate metrology components from vibrating optical elements. A Helium-Neon laser and a position sensor are mounted on one optical bench and a turning flat mirror and a 3-axis fast steering mirror are on the other optical bench. External acoustic disturbances from the speaker are used to generate vibration in the optical components. A speaker housed within an enclosure and a fast steering mirror are the two allowable control actuators and external microphones are used for reference measurement, microphones/accelerometers mounted on or nearby the optical elements and the optical sensor are the three allowable control sensors in the experiment. The generalized predictive controller is used to conjugate the process of the system identification algorithm and the process of the controller design, and can be updated adaptively in the changes of the operating environment. The generalized predictive controllers are designed for each actuator and each sensor: Case 1) Fast steering mirror and position detector, Case 2) Control speaker and position detector, Case 3) Fast steering mirror and external microphones. The generalized predictive controller for all control actuators and sensors is also designed.

The feasibility of a radar instrument working at 95 GHz placed on ISS (International Space Station) to detect very small debris has been investigated and analyzed in this paper. First of all a study about the debris population around the ISS orbit has been taken under consideration by analyzing the debris flux and by determining the preliminary design and mission parameters for Warden instrument, as for instance the pointing angle of the antenna reflector. A technology survey has been also performed to individuate the state of the art in the millimeter wave frequencies band, with particular reference to the transmitter, the front-end, the master oscillator and for A/D converters and DSPs (Digital Signal Processors) suitable for space. The proposed solution for the ISS on board experiment was basically composed of two segments: the on-orbit segment, that is the payload and the ground segment. The on-orbit segment deals with a radar sensor working at 95 GHz which represents a good trade-off between satisfying the limited power consumption available on ISS ExPA (Express Pallet) adapter and having significant range and detection performance necessary for the success of the overall mission, with particular reference to the scientific results. In the overall radar system design major emphasis has been given mainly to the weights, the overall dimensions, and to the EMC problems due to the simultaneous presence of other experiments on the same ExPA. Once fixed the transmitted power, the pulse length and the probability of detection different performances in range are predicting depending on the dimension of the debris and on the approaching speed. For example a detection range (90% PD) of about 180 Km is predicted for a debris of 50 cm running at a relative speed of 10 Km/sec (F=95 GHz, 1 m Cassegrain monopulse antenna, 1000W peak power, 40 /spl mu/sec pulse length, 2.5-100 KHz staggered p.r.f.) Concerning the ground segment, a preliminary architecture has been proposed based upon the extensive use of existing infrastructures and facilities and including a WARDEN control system mainly for all concerns on board payload support, a mission control facility for the general planning of the ground and space resources, and a debris data processing devoted to perform all the activities relevant to the supply of products and services based on payload according to user request. In conclusion the proposed Warden experiment offers several opportunities both in technological field by giving the possibility to verify the sue of millimeter waves components in space environment helpful for future ESA components qualification plan, and in scientific field to validate the concept of a future stand-alone flying on-orbit space-based radar devoted to the detection of very small debris population (0.5 /spl divide/ 1 cm as diameter).

This paper develops a tip-tilt motion controller of fast steering mirror (FSM) in the Giant Magellan telescope (GMT). A mathematical model of tip-tilt motion system of FSM is derived, and then based on this model, stability analysis is carried out. A heuristic adaptive controller is designed for the tip-tilt motion control with modeling error. The heuristic controller consists of two different adaptations such as initial adaptive control and adaptation at steady state errors. Through numerical simulations, the validity of the controller is illustrated. After that this paper addresses several practical issues like disturbance from wind, actuator saturation and resonance frequency from mechanical structure in implementing the controller to the actual system.

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.

Laser communications can enable more efficient and higher bandwidth communications across longer distances than conventional radio frequency (RF) systems. However, beam divergence angles for laser systems are narrower than typical RF systems, and require precise pointing, acquisition, and tracking systems to establish and maintain the link. In addition, typical lasercom links are point-to-point, and not capable of multicast or broadcast. Conventional pointing and tracking (PAT) systems use mechanical gimbals or fast-steering mirrors. Mechanical gimbals may not meet the size, weight, and power (SWaP) constraints for small spacecraft, particularly for multiple concurrent spatially diverse beams. Fast-steering mirrors while compact and efficient have limited aperture size, and many would be needed to provide multiple links over a hemisphere. The Miniature Optical Steered Antenna for Intersatellite Communications (MOSAIC) aims to provide nonmechanical pointing and tracking using liquid lenses, allowing a wide field-of-view and support for multiple concurrent links. Initial work with commercially available liquid lenses showed that liquid lenses can be used in a space environment and assessed spatial coverage. In this work, we model a transmitter using three liquid lenses. One on-axis lens provides focusing capability. Two off-axis and perpendicular lenses provide beam steering, with a fisheye lens amplifying the effect. This provides near-hemispherical pointing up to 170 degrees. We investigate beam quality and divergence using a Zemax model and conduct a link analysis dependent on the beam steering angle and rotation angle. A 25 Mbps link with 200 mW transmit power at 1550 nm (optical C band) and 16-ary pulse position modulation (16-PPM) can be maintained up to 28 km separation with 3 dB margin for an Optotune EL-16-40-TC liquid lens. Losses are primarily due to the liquid lenses limiting aperture size to 16 mm. We also consider the impact of diffusers for increasing numerical aperture through a simple ray transfer analysis and experimental results.

The International Space Station (ISS) provides a unique opportunity for demonstrating instrument and other technologies in the space environment. The ISS orbits the Earth at about 370 km altitude and 51.6o inclination. Its ground track passes over about 90% of the Earth’s population and scientifically interesting surface features. At assembly complete, the ISS can externally accommodate experiments at eight locations on the U.S. Truss. Another ten experiments can be accommodated on the Kibo Japanese Experiments Module Exposed Facility (JEF) of which the U.S. is allocated five sites. The European Columbus Orbital Facility can also accommodate four unpressurized experiments with two allocated for U.S. research use. These locations provide excellent views of either Earth (nadir-mounted orientation) or space (zenith-mounted orientation). This paper will address potential applications of ISS as a platform for demonstrating technologies for future Earth and space science missions.

This paper contains a presentation of the techniques needed to develop simulations for large scale systems which are composed of control systems, optics, and structures. A case study is presented which consists of a generic, controlled specraft whose purpose is to maintain a precise line-of-sight (LOS) in the presence of disturbances and structural deformations. The spacecraft model consists of a six degree-of-freedom (DOF) rigid body with superimposed small structural deformaüons obtained from high-order finite element model. The optical system consists of a telescope with a controlled Fast Steering Mirror (FSM) for maintaining an accurate line-of-sight. Reaction wheels and an Inertial Measurement Unit (IMU) are used for spacecraft attitude control. Disturbances consist of onboard payload motions, sensor noises and reation wheel imbalance. Results of the simulation are shown.

ASTERIA (Arcsecond Space Telescope Enabling Research in Astrophysics) was a CubeSat space telescope that operated in low-Earth orbit, having been deployed from the International Space Station in 2017. A key enabling feature of the ASTERIA mission is the low level of pointing error that has been demonstrated under ASTERIA's fine pointing control mode. Prior to launch, the ASTERIA mission performed analysis and simulations to estimate the in-flight pointing performance. This analysis used reaction wheel models provided by JPL's Small Satellite Dynamics Testbed (SSDT), which had performed Kistler table testing to characterize the jitter caused by the reaction wheels of the Blue Canyon Technologies XACT attitude control unit within ASTERIA. These models were subsequently incorporated into the SSDT simulation. The main source of jitter on ASTERIA is the set of rotating reaction wheels that are used for maintaining fine pointing during observations. Therefore, to benefit all SmallSat missions devoted to photometric or spectroscopic astrophysics applications, an on-orbit flight data acquisition and jitter testing campaign was performed in an attempt to validate the SSDT's simulation models. This paper describes the process used to validate several simulation models and attempt to characterize the ASTERIA jitter by analyzing the size of the resulting spot size for the target stars as a function of wheel speed, having commanded four wheel speeds for each of three stars of known brightness. After summarizing the on-orbit observation and jitter level measurement process, this paper compares the results obtained from the flight operational environment to the corresponding set of simulation jitter levels. Though the flight data obtained during this experiment was not sufficient to validate the simulation's jitter model, other models have been validated with flight data, including the magnetic field, orbit propagation, sun position, and long-duration orbit decay models, which may be used by all other projects that use the SSDT's simulation to increase system capability knowledge.

NASA is leading the design and development of a next-generation CO2 removal system, the four bed carbon dioxide scrubber (4BCO2), and intends to use the International Space Station (ISS) as its testbed. A key component of the system is the blower that provides the airflow through the CO2 sorbent beds. To improve performance and reliability, magnetic levitation (magnetic bearings) will be used in lieu of more conventional bearings (e.g., ball bearings or air bearings) to improve resistance to contaminants and enable extensibility with regards to blower speed, pressure rise and mass flow rate. The blower will pull air from the ISS through an adsorbing desiccant bed and push it through a CO2 sorbent bed and desorbing desiccant bed. The 4BCO2 blower features an overhung permanent magnet motor, a centrally located five-axis, active magnetic bearing system, backup bearings, and an overhung centrifugal impeller in a very compact package. Magnetic bearings are a natural choice for this application due to low power consumption, low transmitted vibration and oil free operation. This article describes the design considerations and design selections for the blower system with a focus on the magnetic bearings. Magnetic FEA of the actuator/sensor system, rotordynamics/controls analysis, and backup bearing drop simulations are discussed in detail. It is expected that the successful implementation of magnetic bearings for this space application will encourage the more widespread adoption in other space applications (e.g., fluid pumps, reaction wheels) that challenge conventional bearing technologies.

Miniaturized attitude control system for nanosatellites