SDO Onboard Ephemeris Generation

The Solar Dynamics Observatory (SDO) spacecraft is a sun-pointing, semi-autonomous satellite that will allow nearly continuous observations of the Sun with a continuous science data downlink. The science requirements for this mission necessitate very strict sun-pointing requirements, as well as continuous ground station connectivity through high gain antennas (HGAs). For SDO s onboard attitude control system to successfully point the satellite at the Sun and the HGAs at the ground stations with the desired accuracy, in addition to the need for accurate sensors it must have good onboard knowledge of the ephemerides of the Sun, the spacecraft, and the ground station. This paper describes the minimum force models necessary for onboard ephemeris generation in support of an attitude control system. The forces that were considered include the Sun s point mass, Moon s point mass, solar radiation pressure (SRP), and the Earth s gravity with varying degree and order of terms of the geopotential

Verification of the Solar Dynamics Observatory High Gain Antenna Pointing Algorithm Using Flight Data

The Solar Dynamics Observatory (SDO) is a NASA spacecraft designed to study the Sun. It was launched on February 11, 2010 into a geosynchronous orbit, and uses a suite of attitude sensors and actuators to finely point the spacecraft at the Sun. SDO has three science instruments: the Atmospheric Imaging Assembly (AIA), the Helioseismic and Magnetic Imager (HMI), and the Extreme Ultraviolet Variability Experiment (EVE). SDO uses two High Gain Antennas (HGAs) to send science data to a dedicated ground station in White Sands, New Mexico. In order to meet the science data capture budget, the HGAs must be able to transmit data to the ground for a very large percentage of the time. Each HGA is a dual-axis antenna driven by stepper motors. Both antennas transmit data at all times, but only a single antenna is required in order to meet the transmission rate requirement. For portions of the year, one antenna or the other has an unobstructed view of the White Sands ground station. During other periods, however, the view from both antennas to the Earth is blocked for different portions of the day. During these times of blockage, the two HGAs take turns pointing to White Sands, with the other antenna pointing out to space. The HGAs handover White Sands transmission responsibilities to the unblocked antenna. There are two handover seasons per year, each lasting about 72 days, where the antennas hand off control every twelve hours. The non-tracking antenna slews back to the ground station by following a ground commanded trajectory and arrives approximately 5 minutes before the formerly tracking antenna slews away to point out into space. The SDO Attitude Control System (ACS) runs at 5 Hz, and the HGA Gimbal Control Electronics (GCE) run at 200 Hz. There are 40 opportunities for the gimbals to step each ACS cycle, with a hardware limitation of no more than one step every three GCE cycles. The ACS calculates the desired gimbal motion for tracking the ground station or for slewing, and sends the command to the GCE at 5 Hz. This command contains the number of gimbals steps for that ACS cycle, the direction of motion, the spacing of the steps, and the delay before taking the first step. The AIA and HMI instruments are sensitive to spacecraft jitter. Pre-flight analysis showed that jitter from the motion of the HGAs was a cause of concern. Three jitter mitigation techniques were developed to overcome the effects of jitter from different sources. The first method is the random step delay, which avoids gimbal steps hitting a cadence on a jitter-critical mode by pseudo-randomly delaying the first gimbal step in an ACS cycle. The second method of jitter mitigation is stagger stepping, which forbids the two antennas from taking steps during the same ACS cycle in order to avoid constructively adding jitter from two antennas. The third method is the inclusion of an instrument No Step Request (NSR), which allows the instruments to request a stoppage in gimbal stepping during the times when they are taking images. During the commissioning phase of the mission, a jitter test was performed onboard the spacecraft. Various sources of jitter, such as the reaction wheels, the High Gain Antenna motors, and the motion of the instrument filter wheels, were examined to determine the level of their effect on the instruments. During the HGA portion of the test, the jitter amplitudes from the single step of a gimbal were examined, as well as the amplitudes due to the execution of various gimbal rates. These jitter levels are compared with the gimbal jitter allocations for each instrument. Additionally, the jitter test provided insight into a readback delay that exists with the GCE. Pre-flight analysis suggested that gimbal steps scheduled to occur during the later portion of an ACS cycle would not be read during that cycle, resulting in a delay in the telemetered current gimbal position. Flight data from the jitter test confirmed this expectation. Analysis is presentehat shows the readback delay does not have a negative impact on gimbal control. The decision was made to consider implementing two of the jitter mitigation techniques on board the spacecraft: stagger stepping and the NSR. Flight data from two sets of handovers, one set without jitter mitigation and the other with mitigation enabled, were examined. The trajectory of the predicted handover was compared with the measured trajectory for the two cases, showing that tracking was not negatively impacted with the addition of the jitter mitigation techniques. Additionally, the individual gimbal steps were examined, and it was confirmed that the stagger stepping and NSRs worked as designed. An Image Quality Test was performed to determine the amount of cumulative jitter from the reaction wheels, HGAs, and instruments during various combinations of typical operations. In this paper, the flight results are examined from a test where the HGAs are following the path of a nominal handover with stagger stepping on and HMI NSRs enabled. In this case, the reaction wheels are moving at low speed and the instruments are taking pictures in their standard sequence. The flight data shows the level of jitter that the instruments see when their shutters are open. The HGA-induced jitter is well within the jitter requirement when the stagger step and NSR mitigation options are enabled. The SDO HGA pointing algorithm was designed to achieve nominal antenna pointing at the ground station, perform slews during handover season, and provide three HGA-induced jitter mitigation options without compromising pointing objectives. During the commissioning phase, flight data sets were collected to verify the HGA pointing algorithm and demonstrate its jitter mitigation capabilities

Reaction Wheel Disturbance Modeling, Jitter Analysis, and Validation Tests for Solar Dynamics Observatory

The Solar Dynamics Observatory (SDO) aims to study the Sun's influence on the Earth by understanding the source, storage, and release of the solar energy, and the interior structure of the Sun. During science observations, the jitter stability at the instrument focal plane must be maintained to less than a fraction of an arcsecond for two of the SDO instruments. To meet these stringent requirements, a significant amount of analysis and test effort has been devoted to predicting the jitter induced from various disturbance sources. One of the largest disturbance sources onboard is the reaction wheel. This paper presents the SDO approach on reaction wheel disturbance modeling and jitter analysis. It describes the verification and calibration of the disturbance model, and ground tests performed for validating the reaction wheel jitter analysis. To mitigate the reaction wheel disturbance effects, the wheels will be limited to operate at low wheel speeds based on the current analysis. An on-orbit jitter test algorithm is also presented in the paper which will identify the true wheel speed limits in order to ensure that the wheel jitter requirements are met

Stochastic performance analysis and staged control system designs for space based interferometers

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.Includes bibliographical references (p. 219-224).Space based interferometry missions have the potential to revolutionize astrometry, providing observations of unprecedented accuracy. Realizing the full potential of these interferometers poses several significant technological challenges. One of the most significant challenges is regulating the light pathlength, from the collecting telescopes to the combining instrument, with nanometer accuracy, despite the presence of vibration induced by internal and external disturbance sources. Due to the wide range of disturbances that act on the optical instruments, a single actuator with simultaneously large control authority and high bandwidth would be necessary to meet the stabilization requirements. Unfortunately no single actuator can meet these requirements. Therefore, a suite of actuators with overlapping strokes and bandwidths is employed; such a construction is termed a "staged actuation system." The objective in the thesis is to develop "staging control" strategies that specify how to utilize the individual actuators in a staged system to satisfy the stabilization requirements. The first task in the staging control design process is to evaluate how actuator capabilities and constraints affect the system performance. Two analytical techniques, based on stochastic Lyapunov and stochastic linearization methods, are utilized to predict the steady-state, closed-loop performance in the presence of actuator nonlinearities such as saturation and quantization. These nonlinearities can severely restrict the achievable performance, and careful consideration of their effects is vitally important for staged controller designs.(cont.) Using this performance prediction methodology, a control synthesis framework is developed which extends [Eta]-optimization techniques by incorporating the effects of actuator nonlinearities. The newly developed framework is then proposed as a formal synthesis tool for staging controller designs. The proposed technique estimates and can directly shape the probability of saturation of each actuator, and determines optimal "hand-offs" of control authority between the actuators. Due to the [Eta]-setting, the controller designs arise directly as solutions of the corresponding necessary conditions, allowing system design studies to be performed easily and quickly. The proposed staged controllers have been demonstrated to achieve the nanometer level stabilization requirements under the expected disturbance environments for space interferometers. The new synthesis technique is also used to analytically quantify the achievable performance and the sensitivity of a particular staging configuration to the individual actuator parameters (size and bandwidths).by Kuo-Chia Liu.Ph.D

A Jitter-Mitigating High Gain Antenna Pointing Algorithm for the Solar Dynamics Observatory

This paper details a High Gain Antenna (HGA) pointing algorithm which mitigates jitter during the motion of the antennas on the Solar Dynamics Observatory (SDO) spacecraft. SDO has two HGAs which point towards the Earth and send data to a ground station at a high rate. These antennas are required to track the ground station during the spacecraft Inertial and Science modes, which include periods of inertial Sunpointing as well as calibration slews. The HGAs also experience handoff seasons, where the antennas trade off between pointing at the ground station and pointing away from the Earth. The science instruments on SDO require fine Sun pointing and have a very low jitter tolerance. Analysis showed that the nominal tracking and slewing motions of the antennas cause enough jitter to exceed the HGA portion of the jitter budget. The HGA pointing control algorithm was expanded from its original form as a means to mitigate the jitter

Investigating On-Orbit Attitude Determination Anomalies for the Solar Dynamics Observatory Mission

The Solar Dynamics Observatory (SDO) was launched on February 11, 2010 from Kennedy Space Center on an Atlas V launch vehicle into a geosynchronous transfer orbit. SDO carries a suite of three scientific instruments, whose observations are intended to promote a more complete understanding of the Sun and its effects on the Earth's environment. After a successful launch, separation, and initial Sun acquisition, the launch and flight operations teams dove into a commissioning campaign that included, among other things, checkout and calibration of the fine attitude sensors and checkout of the Kalman filter (KF) and the spacecraft s inertial pointing and science control modes. In addition, initial calibration of the science instruments was also accomplished. During that process of KF and controller checkout, several interesting observations were noticed and investigated. The SDO fine attitude sensors consist of one Adcole Digital Sun Sensor (DSS), two Galileo Avionica (GA) quaternion-output Star Trackers (STs), and three Kearfott Two-Axis Rate Assemblies (hereafter called inertial reference units, or IRUs). Initial checkout of the fine attitude sensors indicated that all sensors appeared to be functioning properly. Initial calibration maneuvers were planned and executed to update scale factors, drift rate biases, and alignments of the IRUs. After updating the IRU parameters, the KF was initialized and quickly reached convergence. Over the next few hours, it became apparent that there was an oscillation in the sensor residuals and the KF estimation of the IRU bias. A concentrated investigation ensued to determine the cause of the oscillations, their effect on mission requirements, and how to mitigate them. The ensuing analysis determined that the oscillations seen were, in fact, due to an oscillation in the IRU biases. The low frequencies of the oscillations passed through the KF, were well within the controller bandwidth, and therefore the spacecraft was actually following the oscillating biases, resulting in movement of the spacecraft on the order of plus or minus 20 arcsec. Though this level of error met the ACS attitude knowledge requirement of [35, 70, 70] arcsec, 3 sigma, the desire of the ACS and instrument teams was to remove as much of the oscillation as possible. The Kearfott IRUs have an internal temperature controller, designed to maintain the IRU temperature at a constant temperature of approximately 70 C, thus minimizing the change in the bias drift and scale factors of the mechanical gyros. During ground testing of the observatory, it was discovered that the 83-Hz control cycle of the IRU heaters put a tremendous amount of stress on the spacecraft battery. Analysis by the power systems team indicated that the constant charge/discharge on the battery due to the IRU thermal control cycle could potentially limit the life of the battery. After much analysis, the decision was made not to run the internal IRU heaters. Analysis of on orbit data revealed that the oscillations in the IRU bias had a connection to the temperature of the IRU; changes in IRU temperature resulted in changes in the amplitude and period of the IRU biases. Several mitigating solutions were investigated, the result of which was to tune the KF with larger IRU noise assumptions which allows the KF to follow and correct for the time-varying IRU biases

Passive isolator design for jitter reduction in the Terrestrial Planet Finder Coronagraph

Terrestrial Planet Finder (TPF) is a mission to locate and study extrasolar Earth-like planets. The TPF Coronagraph (TPF-C), planned for launch in the latter half of the next decade, will use a coronagraphic mask and other optics to suppress the light of the nearby star in order to collect visible light from such planets. The required contrast ratio of 5e-11 can only be achieved by maintaining pointing accuracy to 4 milli-arcseconds, and limiting optics jitter to below 5 nm. Numerous mechanical disturbances act to induce jitter. This paper concentrates on passive isolation techniques to minimize the optical degradation introduced by disturbance sources. A passive isolation system, using compliant mounts placed at an energy bottleneck to reduce energy transmission above a certain frequency, is a low risk, flight proven design approach. However, the attenuation is limited, compared to an active system, so the feasibility of the design must be demonstrated by analysis. The paper presents the jitter analysis for the baseline TPF design, using a passive isolation system. The analysis model representing the dynamics of the spacecraft and telescope is described, with emphasis on passive isolator modeling. Pointing and deformation metrics, consistent with the TPF-C error budget, are derived. Jitter prediction methodology and results are presented. Then an analysis of the critical design parameters that drive the TPF-C jitter response is performed

Wide-Field Infrared Survey Telescope (WFIRST) Integrated Modeling

Contents: introduction to WFIRST (Wide-Field Infrared Survey Telescope) and integrated modeling; WFIRST stability requirement summary; instability mitigation strategies; dynamic jitter results; STOP (structural-thermal-optical performance) (thermal distortion) results; STOP and jitter capability limitations; model validation philosophy