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The Cassini spacecraft was launched on October 15, 1997. It uses thrusters to perform many spacecraft control functions: to detumble the spacecraft, to maintain three-axis attitude control, to perform small trajectory correction burns, to desaturate the reaction wheels, and others.
The Clementine Mission is summarized from purposes of the mission through details of the instruments. The mission was a joint Department of Defense and NASA mission to use lighter instruments and other equipment developed by the military for space exploration. New lightweight items on Clementine include batteries, optical sensors, reaction wheels, star trackers, release systems, and inertial measurement units. Other items include a compression chip set, a solid state data recorder, autonomous navigation, and a 32-bit RISC computer.
On 5 September 2012, Dawn concluded its successful exploration of Vesta, the second most massive object in the main asteroid belt. The spacecraft departed after 14 months in orbit and is now using its ion propulsion system to travel to dwarf planet Ceres, the most massive main-belt asteroid. The principal activity now is thrusting with the ion propulsion system to provide the 3.3 km/s required to rendezvous with Ceres early in 2015. Because two of the four reaction wheels have experienced faults and are likely unrecoverable, a substantial effort has been invested in preparing for Ceres operations with alternate attitude control methods. The project has engaged in an intensive campaign to reduce hydrazine expenditures, which has resulted in a significant increase in the hydrazine expected to be available for Ceres. Based on this work, studies provide good confidence that the required activities at Ceres can be completed. This paper describes post-Vesta operations, including measures taken to conserve hydrazine as well as other preparations for Ceres.
To maintain the required WFIRST Coronagraph starlight suppression performance in a realistic space environment, a low order wavefront sensing and control (LOWFS/C) subsystem is necessary. The LOWFS/C uses the rejected stellar light from coronagraph to sense and suppress the telescope pointing drift and jitter as well as the low order wavefront errors due to changes in thermal loading on the telescope and the rest of the observatory. In this paper we will present an overview of the low order wavefront sensing and control subsystem for the WFIRST Coronagraph and describe the WFIRST Coronagraph LOWFS function, its design, and modeled performance. We will present experimental results on a dedicated LOWFS/C testbed that show that the LOWFS/C subsystem not only can sense pointing errors better than 0.2 mas but has also experimentally demonstrated closed loop pointing error suppression with residuals better than 0.4 mas rms per axis for the vast majority of observatory reaction wheel speeds.
Dawn has conducted an extensive exploration of dwarf planet Ceres, the largest object between the Sun and Pluto that had not previously been visited by a spacecraft. Following its arrival at Ceres in March 2015, Dawn acquired all the planned data from four circular polar orbits ranging in altitude from 13,600 km to 385 km. After the successful conclusion of its primary mission in June 2016, Dawn's mission was extended, and new investigations, not previously considered, were conducted from three new orbits. The mission has provided a uniquely detailed view of the first dwarf planet discovered. The overall strategy for exploring Ceres was based strongly on the extremely successful 16 months of Vesta operations. Nevertheless, the loss of two of the spacecraft's four reaction wheels before arrival at Ceres necessitated some important changes in order to conserve hydrazine. These changes were so effective that Dawn has been able to operate beyond the expected end of life, acquiring significantly more data than planned. This paper describes Ceres operations in the primary and extended missions as well as some of the major findings there.
As it has for the past few years, numerical modeling is being used to predict the on-orbit, high-contrast imaging performance of the WFIRST coronagraph, which was recently defined to be solely a technology demonstrator. A consequence has been a realignment of modeling priorities and revised applications of modeling uncertainty factors and margins, which apply to multiple factors such as pointing and wavefront jitter, thermally-induced deformations, polarization, and aberration sensitivities. At the same time, the models have increased in fidelity as additional parameters have been added, such as time-dependent pupil shear and mid-spatial-frequency deformations of the primary and secondary mirrors, detector effects, and reaction-wheel-speed-dependent pointing and wavefront jitter.
A small thruster, that will herein be called the Hydrazine milliNewton Thruster (HmNT), has been designed and vacuum hot-fire tested, demonstrating a range of thrust and impulse far below the current state-of-the-art (SOA) hydrazine thruster. Steady-state thrust levels of 35–135 mN, and minimum impulse-bits of 25–120 μN-s were demonstrated over a typical operating range of inlet pressures. These capabilities provide new spacecraft and mission design options, including the following: Functional backup for Reaction Wheel Assemblies (RWAs), which can be used to prolong the useful life of RWAs; RWA replacement for certain missions to save mass, power, and cost; Precision delta-V that can be used for: proximity operations and docking maneuvers, formation flying, primary attitude control for SmallSats and CubeSats, and primary delta-V for CubeSats.
Dawn has been in orbit around Ceres since March 2015. In 2016, Dawn completed its prime mission of exploring Vesta and Ceres, and it completed an extended mission at the dwarf planet in 2017. The spacecraft has lost all reaction wheel control, making attitude control dependent on a dwindling supply of hydrazine. Nevertheless, in 2017 the project was approved for a second and final extended mission at Ceres. Starting in April 2018, Dawn used its ion propulsion system to maneuver to two new orbits. These highly elliptical orbits provided hydrazine-efficient opportunities for the acquisition of valuable new data. Now in the second of these orbits, Dawn's lowest altitude is only 9% of the lowest altitude previously in the mission, enabling significant improvements in the spatial resolution of the data. We describe Dawn operations in its second extended mission, the nature of the orbits, and the observation campaigns. We also summarize expectations for the end of the mission.
Explosions, collisions, and other catastrophic breakups of launch vehicle orbital stages and satellites continue to be major contributors to the generation of orbital debris. Both launch vehicles and payload satellites typically have several types of stored energy sources on board, any of which might result in energetic breakups and the creation of debris after their mission has ended. These energy sources include propulsion systems, pressure vessels, reaction wheels, control moment gyros, heat pipes, and power systems. NASA, ESA, JAXA and other space-faring organizations have requirements in place to limit the growth of the orbital debris population by passivating space vehicles that remain in orbit after their missions have ended. In this paper, we review current spacecraft passivation philosophies and principles, as well as how those principles have been applied in practice. In particular, we focus on how NASA programs have addressed spacecraft passivation. We begin by considering and reviewing general passivation requirements, with specific emphasis on pressure vessel passivation. We discuss passivation approaches used in several recent NASA missions as well as some practical considerations in spacecraft passivation, and conclude by providing some summary guidelines regarding what may be considered acceptable (reduced) pressure level targets (depending on the tank commodity and the type of propulsion system) that could allow the pressure vessel to be considered in a passivated state.
The CloudSat satellite was launched from Vandenberg Air Force Base, in California, on April 28, 2006. It was launched with the CALIPSO satellite, a joint mission between the National Aeronautics and Space Administration (NASA) and the Centre National D'Etudes Spatiales (CNES). CloudSat and CALIPSO were designed to fly in formation with each other, to combine their complimentary scientific measurements. The two missions were maneuvered into the Earth Science Afternoon Constellation (A-Train) shortly after launch and have flown in formation together since then. Originally designed to fly for 22 months, the CloudSat satellite has far exceeded expectations. Inflight anomalies on CloudSat necessitated changes to both the daily concept of operations and formation flying operations with CALIPSO. In 2011, a severe battery anomaly restricted CloudSat science data collection and maneuverability to the sunlit portions of the orbit. In 2017, CloudSat suffered a reaction wheel failure, triggering the criteria established to exit the A-Train. CloudSat successfully exited the A-Train in early 2018, suspending formation flying and joint science observations with CALIPSO. The joint CloudSat/CALIPSO science team continued collaboration with discussions focused on resuming joint observations. As both satellites were aging and CALIPSO was nearing the exhaustion of satellite propellant, the decision was made to lower CloudSat to the CALIPSO graveyard orbit, to be joined by CALIPSO as soon as feasible. With NASA’s approval, CloudSat was maneuvered to the lower orbit over the course of several months in mid-2018. Once CloudSat was in place, and with approval from CNES, CALIPSO began the delicate dance of lowering their orbit and re-aligning the satellite ground tracks. CloudSat and CALIPSO resumed formation flying in the newly dubbed C-Train (CloudSat and CALIPSO) orbit in the Fall of 2018. The ingenuity of the CloudSat and CALIPSO teams made the concept of the C-Train a reality. This paper will discuss the decision processes and the operational challenges the CloudSat flight team faced while designing and implementing the C-Train.
Over the past two years, JPL has used the ASTERIA (Arcsecond Space Telescope Enabling Research In Astrophysics) CubeSat as an in-flight test platform during extended missions. ASTERIA successfully completed its prime mission in early 2018, and continued to operate in low Earth orbit (LEO) for an additional twenty months. This paper describes demonstrations that were performed on the spacecraft and on the ground-based testbed during the extended mission. These demonstrations fall into three categories: Autonomy technology maturation, hardware characterization, and science discovery. Autonomy technology maturation supported three development efforts. The first shifted the spacecraft commanding paradigm from time-based sequences to Task Networks (tasknets), which allow simpler commanding and more robust onboard execution. The second demonstrated onboard orbit determination in Low Earth Orbit (LEO) without GPS. This activity used a fully-independent means of spacecraft orbit determination for Earth orbiters using only passive imaging. The third technology provided in situ hardware health state estimation using a model-based reasoning technique. These three technologies were demonstrated either in flight or on the testbed individually, and then were combined to demonstrate the capability to perform autonomous navigation on board without ground intervention, even in the presence of anomalies. Hardware characterization involved both onboard and ground-based activities. On board, nonstandard attitude control modes were commanded to characterize the spacecraft pointing jitter as a function of target brightness, reaction wheel speed, controller gain, and the number of guide stars. The results provide insights into the contribution of jitter to the ASTERIA photometry and inform the feasibility of future astrophysics small satellite missions for which jitter control is an enabling technology. On the ground, the ASTERIA Operations Team coordinated with Amazon Web Services (AWS) to configure their new ground stations to communicate with ASTERIA to prove out their viability. ASTERIA used AWS ground stations for nominal operations for the last four months of the mission. Finally, ASTERIA continued to perform exoplanet science as the spacecraft was well-suited to execute long-term monitoring of stars such as alpha Centauri to search for small transiting planets. The science team also imaged a number of interesting objects including a comet, an asteroid, cities at night, and the moon, and coordinated with other projects on Targets of Opportunity for follow-up confirmations and co-observations. Throughout the prime and the extended missions, the ASTERIA spacecraft proved to be a mighty platform that “will go into history as an innovative milestone.”[1 - Zurbuchen]
Explosions, collisions, and other catastrophic breakups of launch vehicle orbital stages and satellites continue to be major contributors to the generation of orbital debris. Both launch vehicles and payload satellites typically have several types of stored energy sources on board, any of which might result in energetic breakups and the creation of debris after their mission has ended. These energy sources include propulsion systems, pressure vessels, reaction wheels, control moment gyros, heat pipes, and power systems. NASA, ESA, JAXA and other space-faring organizations have requirements in place to limit the growth of the orbital debris population by passivating space vehicles that remain in orbit after their missions have ended. In this paper, we review current spacecraft passivation philosophies and principles, as well as how those principles have been applied in practice. In particular, we focus on how NASA programs have addressed spacecraft passivation. We begin by considering and reviewing general passivation requirements, with specific emphasis on pressure vessel passivation. We then discuss passivation approaches used in several recent NASA missions as well as some practical considerations in spacecraft passivation, and conclude by providing some summary guidelines regarding what may be considered acceptable (reduced) pressure level targets (depending on the tank commodity and the type of propulsion system) that could allow the pressure vessel to be considered in a passivated state.
Solar Cruiser is a small satellite Technology Demonstration Mission (TDM) to mature solar sail propulsion technology using a solar sail larger than 1600 square meters, demonstrating performance as a propulsion system and a stable pointing platform for science observations in an artificial halo orbit sunward of the Sun-Earth Lagrange Point 1 (sub-L1). For the “sailcraft” to meet mission ob- jectives, there are several unique attitude control challenges that the Attitude De- termination and Control System (ADCS) must overcome. Large disturbance tor- ques, primarily due to sail deformations coupled with off-sun pointing angles, make it more difficult to maintain adequate controls performance and manage accumulated momentum on the control actuators. The large-amplitude, low- frequency flexible body modes of the sail, in concert with noisy sensors and ac- tuators, make it challenging to mitigate control-structure interactions and main- tain fine pointing capabilities. Stability and control performance during sail de- ployment is complicated by rapidly and widely varying inertias. The Solar Cruiser ADCS effectively addresses these challenges using a simple, traditional control system – including momentum management actuators that use bang- bang control to constrain internal accumulated momentum, a low-pass controls filter and an attitude determination Kalman Filter (KF) that blends multiple star tracker solutions, and a reaction wheel assembly (RWA) with controller gains tuned to specific configurations or inertias. Solar Cruiser’s attitude control ap- proach, and lessons learned from its development, establishes a state of the art of high value to future solar sail missions and other small spacecraft with large de- ployable structures operating in deep space.
Solar Cruiser is a small satellite Technology Demonstration Mission (TDM) to mature solar sail propulsion technology using a solar sail larger than 1600 square meters, demonstrating performance as a propulsion system and a stable pointing platform for science observations in an artificial halo orbit sunward of the Sun-Earth Lagrange Point 1 (sub-L1). For the “sailcraft” to meet mission objectives, there are several unique attitude control challenges that the Attitude Determination and Control System (ADCS) must overcome. Large disturbance torques, primarily due to sail deformations coupled with off-sun pointing angles, make it more difficult to maintain adequate controls performance and manage accumulated momentum on the control actuators. The large-amplitude, low-frequency flexible body modes of the sail, in concert with noisy sensors and actuators, make it challenging to mitigate control-structure interactions and maintain fine pointing capabilities. Stability and control performance during sail deployment is complicated by rapidly and widely varying inertias. The Solar Cruiser ADCS effectively addresses these challenges using a simple, traditional control system – including momentum management actuators that use bang-bang control to constrain internal accumulated momentum, a low-pass controls filter and an attitude determination Kalman Filter (KF) that blends multiple star tracker solutions, and a reaction wheel assembly (RWA) with controller gains tuned to specific configurations or inertias. Solar Cruiser’s attitude control approach, and lessons learned from its development, establishes a state of the art of high value to future solar sail missions and other small spacecraft with large deployable structures operating in deep space.
Solar Cruiser is a small (ESPA-class) satellite Technology Demonstration Mission (TDM) to mature solar sail propulsion technology using a solar sail larger than 1600 square meters, demonstrating performance both as a propulsion system and a stable pointing platform for science observations in an artificial halo orbit sunward of the Sun-Earth Lagrange Point 1 (sub-L1). To ensure attitude control throughout the mission, momentum accumulated on the reaction wheels (RWs) used for attitude control must be managed such that the sailcraft does not lose control due to RW momentum saturation. Momentum builds up on the wheels from environmental disturbance torques caused by solar radiation pressure combined with a center of mass (CM)/center of pressure (CP) offset, deformed sail shape, and an off-sun pointing angle, plus other factors. Solar Cruiser mitigates this momentum build up by utilizing an Active Mass Translator (AMT) that maintains pitch and yaw momentum by trimming the CM/CP offsets, and thrusters to maintain roll momentum. A survey was conducted by the Solar Cruiser team to assess the feasibility and tradeoffs of novel momentum management concepts such as Reflectivity Control Devices (RCD’s), different thruster configurations, and control vanes and other articulated control surfaces. In addition, techniques to reduce disturbance torque buildup, such as reducing boom tip deflections and clock angle control, were assessed. Similar sailcraft momentum management strategies can be used for future missions such as space weather monitoring and Earth magnetotail science missions
Near Earth Asteroid Scout (NEA Scout) was a mission to test solar sail propulsion for orbital transfer from cislunar space to flyby and image an asteroid. Had it succeeded, one of the mission goals was to characterize the solar torque on the sail to ensure successful attitude control for the orbit transfer and imaging the asteroid. The simulation used to develop the flight attitude control software uses the generalized model for solar sails, a tensor equation of the forces and torques on sails of arbitrary shape. Rios-Reyes and Scheeres developed a general process to update the torque tensor coefficients using estimates of sail torque over a range of directions to the sun. Their process was adapted and implemented for the specific case of NEA Scout using spacecraft telemetry collected during sail characterization maneuvers in combination with simulation models and parameters. The NEA Scout maneuvers were limited to the operating range of the mission and constraints of the control hardware and allowed safe testing of each attitude before proceeding to the next. The NEA Scout reaction wheel speeds are used to measure accumulated momentum, while the Active Mass Translator (AMT) position is used to subtract out the torque from the center of mass crossed with the sail force and isolate the torque from only the sail shape. The process was tested by running attitude control simulations of the characterization maneuvers, generating simulated telemetry, estimating the solar torques, then using a least squares estimating the solar torque coefficients using least-squares and then performing a least-squares fit to the solar torque tensor coefficients. These estimated coefficients were tested by evaluating the solar torques under the same conditions as the simulated telemetry and comparing to the true simulated torques. Solar force model updates can be performed separately by observing the effect of the sail on the trajectory, and the torque model can be refined using those solar force updates. This process met the needs of the NEA Scout mission and can be adapted to characterize the solar torque for other missions with different sails.
Near Earth Asteroid Scout (NEA Scout) was a mission to test solar sail propulsion for orbital transfer from cislunar space to flyby and image an asteroid. One of the goals of the mission was to characterize the solar torque on the sail to ensure successful attitude control for the orbit transfer and imaging the asteroid. The simulation used to develop the flight attitude control software uses the generalized model for solar sails, a tensor equation of the forces and torques on sails of arbitrary shape. Rios-Reyes and Scheeres developed a general process to update the torque tensor coefficients using estimates of sail torque over a range of directions to the sun. Their process was adapted and implemented for the specific case of NEA Scout using spacecraft telemetry collected during sail characterization maneuvers in combination with simulation models and parameters. The NEA Scout maneuvers were limited to the operating range of the mission and constraints of the control hardware and allowed safe testing of each attitude before proceeding to the next. The NEA Scout reaction wheel speeds are used to measure accumulated momentum, while the Active Mass Translator (AMT) position is used to subtract out the torque from the center of mass crossed with the sail force and isolate the torque from only the sail shape. The process was tested by running attitude control simulations of the characterization maneuvers, generating simulated telemetry, estimating the solar torques, then using a least squares estimating the solar torque coefficients using least-squares and then performing a least-squares fit to the solar torque tensor coefficients. These estimated coefficients were tested by evaluating the solar torques under the same conditions as the simulated telemetry and comparing to the true simulated torques. Solar force model updates can be performed separately by observing the effect of the sail on the trajectory, and the torque model can be refined using those solar force updates. This process met the needs of the NEA Scout mission and can be adapted to characterize the solar torque for other missions with different sails.
One of the driving structural requirements of the Habitable Exo-Planet telescope is to maintain line of sight stability between the primary and secondary mirror of ≤ 5 milli-arc seconds. Dynamic analyses of two configurations of a proposed Habitable Exo-Planet 4 meter off-axis telescope structure were performed to predict effects of a vibration input on primary/secondary mirror alignment. The dynamic disturbance used as the forcing function was the James Webb Space Telescope reaction wheel assembly vibration emission specification level. The objective of these analyses was to predict order-of-magnitude performance for various structural configurations which contribute to efforts in defining the Habitable Exo-Planet structural design’s global architecture. Two variations of the basic architectural design were analyzed. Relative motion between the PM and the SM for each design configuration are reported.