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NASA NTRS2023

Earth/Moon System Star Tracking Sensor for Improved Deep Space Laser Communications

The goal of this research is to improve deep space communication links by exploring a new concept in stellar navigation that uses images of the Earth/Moon system to obtain precise pointing information so that the laser signal from the satellite can be directed at the ground station with unprecedented accuracy. Currently, the concept for satellite pointing relies on a dual axis star tracker that compares images of the celestial sphere to a star catalog to obtain a 3-axis orientation solution. Most star trackers are designed with a wide field of view (required to obtain enough bright starts to achieve a navigational fix) and only provide reliable pitch and yaw estimates thus necessitating the use of a second star tracker looking at an axis perpendicular to the first one to obtain an accurate estimate of the roll of the spacecraft. In the new concept the optical design will be modified to ensure that both the Earth and the Moon will be within a narrow field of view centered on Earth relative to the field of view of most star tracking systems. Because of the proximity of the bodies, it is possible to obtain highly accurate estimates of the pitch, yaw and roll of the spacecraft relative to the Earth/Moon system. The new tracker will allow for accurate 3-axis control of the spacecraft from a single tracker thus allowing the second star tracker, which would normally be pointed in a direction perpendicular to the one pointed at earth, to obtain the spacecraft roll estimate. This wide field star tracker facilitates functions like the initial satellite pointing during the acquisition phase of establishing the laser communication link or solving the lost in space problem should the system need to be reset. This new design will be tested in simulation and compared for 3-axis pointing accuracy against a standard commercial star tracker.

Star tracker↗

NASA NTRS2016

Extended Bright Bodies - Flight and Ground Software Challenges on the Cassini Mission at Saturn

Extended bright bodies in the Saturn environment such as Saturn's rings, the planet itself, and Saturn's satellites near the Cassini spacecraft may interfere with the star tracker's ability to find stars. These interferences can create faulty spacecraft attitude knowledge, which would decrease the pointing accuracy or even trip a fault protection response on board the spacecraft. The effects of the extended bright body interference were observed in December of 2000 when Cassini flew by Jupiter. Based on this flight experience and expected star tracker behavior at Saturn, the Cassini AACS operations team defined flight rules to suspend the star tracker during predicted interference windows. The flight rules are also implemented in the existing ground software called Kinematic Predictor Tool to create star identification suspend commands to be uplinked to the spacecraft for future predicted interferences. This paper discusses the details of how extended bright bodies impact Cassini's acquisition of attitude knowledge, how the observed data helped the ground engineers in developing flight rules, and how automated methods are used in the flight and ground software to ensure the spacecraft is continuously operated within these flight rules. This paper also discusses how these established procedures will continue to be used to overcome new bright body challenges that Cassini will encounter during its dips inside the rings of Saturn for its final orbits of a remarkable 20-year mission at Saturn.

Sung, Tina S.↗

NASA NTRS2003

Star Identification Using a Triplet Algorithm

The stars observed by onboard star trackers must be identified in order for the star tracker information to be used for attitude determination. Star identification associates the observed stars in the body frame with catalog stars, the positions of which in an inertial frame are well known. If the attitude is approximately known (from less accurate sensors) identification is relatively simple, but if there is no prior attitude knowledge (the Lost in Space problem) star identification is quite difficult. A new, reliable, and accurate star identification algorithm has been developed for use in a batch, ground-based attitude determination system. The identification algorithm uses input from one or more star trackers, each of which can observe three or more stars simultaneously. If observations are available at different times, rate data is used to relate them. The algorithm includes the following steps that are automatically performed: 1) Observations are examined to determine the optimal set of simultaneous star observations in a single tracker. 2) The selected optimal observations are identified using a triplet match. 3) The identified stars in the selected tracker at the selected time are used to compute an attitude at that time. 4) The rate data is refined by determination of biases that minimize the dispersion of repeated observations of the same stars. 5) The refined rate data is used to propagate the attitude to the times of each star observation. 6) The attitude at each time is used to associate all stars observed at that time, in each tracker, with catalog stars. The overall algorithm is very dependable at producing accurate star identification with no initial attitude input.

Hashmall, Joseph A.↗

NASA NTRS1986

CSTAR star catalogue development

The Continuous Stellar Tracking Attitude Reference (CSTAR) system is an in-house project for the Space Station to provide high accuracy, drift free attitude and angular rate information for the GN&C system. Constraints exist on the star catalogue incorporated in the system. These constraints include the following: mass memory allocated for catalogue storage, star tracker imaging sensitivity, the minimum resolvable separation angle between stars, the width of the field of view of the star tracker, and the desired number of stars to be tracked in a field of view. The Smithsonian Astrophysical Observatory (SAO) catalogue is the basis reference for this study. As it stands, the SAO does not meet the requirements of any of the above constraints. Star selection algorithms have been devised for catalogue optimization. Star distribution statistics have been obtained to aid in the development of these rules. VAX based software has been developed to implement the star selection algorithms. The software is modular and provides a design tool to tailor the catalogue to available star tracker technology. The SAO catalogue has been optimized for the requirements of the present CSTAR system.

Uhde-Lacovara, J. A.↗

NASA NTRS2009

Necessary and Sufficient Conditions for Attitude Estimation in Fractionated Spacecraft Systems

This paper addresses the problem of attitude estimation in fractionated spacecraft clusters. Each module in the cluster may have either a star-tracker, a relative attitude sensor, or both. Using results in nonlinear ob- servability theory, we provide graph-theoretic sufficient conditions for the attitude of every module to be observable. In particular we show that the attitude of every module in the cluster can be observed if every module has either a star tracker with non-collinear stars, or there is a path through the sensing network from a module with a star tracker to the module without a star tracker, and each of the relative measurements along the path has either multiple non-collinear beacons or a single beacon that is not parallel to the rotation vector of the target module.

Blackmore, Lars↗

NASA NTRS2011

Backup Alignment Devices on Shuttle: Heads-Up Display or Crew Optical Alignment Sight

NASA s Space Shuttle was built to withstand multiple failures while still keeping the crew and vehicle safe. Although the design of the Space Shuttle had a great deal of redundancy built into each system, there were often additional ways to keep systems in the best configuration if a failure were to occur. One such method was to use select pieces of hardware in a way for which they were not primarily intended. The primary function of the Heads-Up Display (HUD) was to provide the crew with a display of flight critical information during the entry phase. The primary function of the Crew Optical Alignment Sight (COAS) was to provide the crew an optical alignment capability for rendezvous and docking phases. An alignment device was required to keep the Inertial Measurement Units (IMUs) well aligned for a safe Entry; nominally this alignment device would be the two on-board Star Trackers. However, in the event of a Star Tracker failure, the HUD or COAS could also be used as a backup alignment device, but only if the device had been calibrated beforehand. Once the HUD or COAS was calibrated and verified then it was considered an adequate backup to the Star Trackers for entry IMU alignment. There were procedures in place and the astronauts were trained on how to accurately calibrate the HUD or COAS and how to use them as an alignment device. The calibration procedure for the HUD and COAS had been performed on many Shuttle missions. Many of the first calibrations performed were for data gathering purposes to determine which device was more accurate as a backup alignment device, HUD or COAS. Once this was determined, the following missions would frequently calibrate the HUD in order to be one step closer to having the device ready in case it was needed as a backup alignment device.

Chavez, Melissa A.↗

NASA NTRS1995

Development of a Robust star identification technique for use in attitude determination of the ACE spacecraft

The Advanced Composition Explorer (ACE) spacecraft is designed to fly in a spin-stabilized attitude. The spacecraft will carry two attitude sensors - a digital fine Sun sensor and a charge coupled device (CCD) star tracker - to allow ground-based determination of the spacecraft attitude and spin rate. Part of the processing that must be performed on the CCD star tracker data is the star identification. Star data received from the spacecraft must be matched with star information in the SKYMAP catalog to determine exactly which stars the sensor is tracking. This information, along with the Sun vector measured by the Sun sensor, is used to determine the spacecraft attitude. Several existing star identification (star ID) systems were examined to determine whether they could be modified for use on the ACE mission. Star ID systems which exist for three-axis stabilized spacecraft tend to be complex in nature and many require fairly good knowledge of the spacecraft attitude, making their use for ACE excessive. Star ID systems used for spinners carrying traditional slit star sensors would have to be modified to model the CCD star tracker. The ACE star ID algorithm must also be robust, in that it will be able to correctly identify stars even though the attitude is not known to a high degree of accuracy, and must be very efficient to allow real-time star identification. The paper presents the star ID algorithm that was developed for ACE. Results from prototype testing are also presented to demonstrate the efficiency, accuracy, and robustness of the algorithm.

Woodard, Mark↗

NASA NTRS2022

Modeling Meteoroid Impacts on the Juno Spacecraft

Events which meet certain criteria from star-tracker images on board the Juno spacecraft have been proposed to be due to interplanetary dust particle impacts on its solar arrays. These events have been suggested to be caused by particles with diameters larger than 10 μm. Here, we compare the reported event rates to expected dust-impact rates using dynamical meteoroid models for the four most abundant meteoroid/dust populations in the inner solar system. We find that the dust-impact rates predicted by dynamical meteoroid models are not compatible with either the Juno observations in terms of the number of star-tracker events per day, or with the variations of dust flux on Juno's solar panels with time and position in the solar system. For example, the rate of star-tracker events on Juno's antisunward surfaces is the largest during a period in which Juno is expected to experience the peak impact fluxes on the opposite, sunward hemisphere. We also investigate the hypothesis of dust leaving the Martian Hill sphere originating either from the surface of Mars itself or from one of its moons. We do not find such a hypothetical source to be able to reproduce the star-tracker event-rate variations observed by Juno. We conclude that the star-tracker events observed by Juno are unlikely to be the result of instantaneous impacts from the zodiacal cloud.

Interplanetary dust↗

NASA NTRS2019

Beaconless Optical Communication System Constraints

Deep-space optical communication will enable increased science return and public engagement for robotic and manned missions. The IROC project is studying a beaconless optical communication system for Mars data downlink. A star tracker provides the optical communications pointing information in place of an uplink targeting beacon. The configuration presented in this paper includes a star tracker that is aligned co-boresighted with the optical communication axis. This co-boresight configuration was not discussed in prior work, as it was assumed that large Sun-Probe-Earth keep-out angle requirements for operation of the star tracker would cause significant communication outages. In this paper it is shown that the use of an optimal mechanical mounting angle combined with an advanced star tracker has the capability to yield up to 92% communication availability for an example five-year Mars mission.

Aretskin-Hariton, Eliot D.↗

NASA NTRS1976

Advanced application flight experiments precision attitude determination system. Volume 2: System tests

The performance capability of each of two precision attitude determination systems (PADS), one using a strapdown star tracker, and the other using a single-axis gimbal star tracker was measured in the laboratory under simulated orbit conditions. The primary focus of the evaluation was on the contribution to the total system accuracy by the star trackers, and the effectiveness of the software algorithms in functioning with actual sensor signals. A brief description of PADS, the laboratory test configuration and the test facility, is given along with a discussion of the data handling and display, laboratory computer programs, PADS performance evaluation programs, and the strapdown and gimbal system tests. Results are presented and discussed.

Source record↗

NASA NTRS1977

Performance tests of two precision attitude determination systems

Results of laboratory performance tests of two satellite attitude determination systems are given. One system employed a strapdown star tracker and gyro assembly, the other a single axis, gimballed star tracker and a gyro assembly. The laboratory tests simulated those orbit conditions which would be experienced on a three axis stabilized, earth pointed satellite in geosynchronous orbit. A ground-fixed laboratory test was performed in which system axes remained stationary in the laboratory coordinates while revolving star beams stimulated the star trackers. The laboratory instrumentation techniques used to meet the stringent accuracy requirements are described. Results are presented which show both systems met the performance goal of 3.6 arc seconds. Comparative analyses of both systems are also discussed.

Mcaloon, K. J.↗

NASA NTRS2010

Real-Time Visualization of Spacecraft Telemetry for the GLAST and LRO Missions

GlastCam and LROCam are closely-related tools developed at NASA Goddard Space Flight Center for real-time visualization of spacecraft telemetry, developed for the Gamma-Ray Large Area Space Telescope (GLAST) and Lunar Reconnaissance Orbiter (LRO) missions, respectively. Derived from a common simulation tool, they use related but different architectures to ingest real-time spacecraft telemetry and ground predicted ephemerides, and to compute and display features of special interest to each mission in its operational environment. We describe the architectures of GlastCam and LROCam, the customizations required to fit into the mission operations environment, and the features that were found to be especially useful in early operations for their respective missions. Both tools have a primary window depicting a three-dimensional Cam view of the spacecraft that may be freely manipulated by the user. The scene is augmented with fields of view, pointing constraints, and other features which enhance situational awareness. Each tool also has another "Map" window showing the spacecraft's groundtrack projected onto a map of the Earth or Moon, along with useful features such as the Sun, eclipse regions, and TDRS satellite locations. Additional windows support specialized checkout tasks. One such window shows the star tracker fields of view, with tracking window locations and the mission star catalog. This view was instrumental for GLAST in quickly resolving a star tracker mounting polarity issue; visualization made the 180-deg mismatch immediately obvious. Full access to GlastCam's source code also made possible a rapid coarse star tracker mounting calibration with some on the fly code adjustments; adding a fine grid to measure alignment offsets, and introducing a calibration quaternion which could be adjusted within GlastCam without perturbing the flight parameters. This calibration, from concept to completion, took less than half an hour. Both GlastCam and LROCam were developed in the C language, with non-proprietary support libraries, for ease of customization and portability. This no-blackboxes aspect enables engineers to adapt quickly to unforeseen circumstances in the intense operations environment. GlastCam and LROCam were installed on multiple workstations in the operations support rooms, allowing independent use by multiple subsystems, systems engineers and managers, with negligible draw on telemetry system resources.

Stoneking, Eric T.↗

NASA NTRS1993

The Spacelab IPS Star Simulator

The cost of doing business in space is very high. If errors occur while in orbit the costs grow and desired scientific data may be corrupted or even lost. The Spacelab Instrument Pointing System (IPS) Star Simulator is a unique test bed that allows star trackers to interface with simulated stars in a laboratory before going into orbit. This hardware-in-the-loop testing of equipment on earth increases the probability of success while in space. The IPS Star Simulator provides three fields of view 2.55 x 2.55 deg each for input into star trackers. The fields of view are produced on three separate monitors. Each monitor has 4096 x 4096 addressable points and can display 50 stars (pixels) maximum at a given time. The pixel refresh rate is 1000 Hz. The spectral output is approximately 550 nm. The available relative visual magnitude range is two to eight visual magnitudes. The star size is less than 100 arcsec. The minimum star movement is less than 5 arcsec and the relative position accuracy is approximately 40 arcsec. The purpose of this paper is to describe the IPS Star Simulator design and to provide an operational scenario so others may gain from the approach and possible use of the system.

Wessling, Francis C., III↗

NASA NTRS1993

The Spacelab IPS Star Simulator

The cost of doing business in space is very high. If errors occur while in orbit the costs grow and desired scientific data may be corrupted or even lost. The Spacelab Instrument Pointing System (IPS) Star Simulator is a unique test bed that allows star trackers to interface with simulated stars in a laboratory before going into orbit. This hardware-in-the loop testing of equipment on earth increases the probability of success while in space. The IPS Star Simulator provides three fields of view 2.55 x 2.55 degrees each for input into star trackers. The fields of view are produced on three separate monitors. Each monitor has 4096 x 4096 addressable points and can display 50 stars (pixels) maximum at a given time. The pixel refresh rate is 1000 Hz. The spectral output is approximately 550 nm. The available relative visual magnitude range is 2 to 8 visual magnitudes. The star size is less than 100 arc seconds. The minimum star movement is less than 5 arc seconds and the relative position accuracy is approximately 40 arc seconds. The purpose of this paper is to describe the LPS Star Simulator design and to provide an operational scenario so others may gain from the approach and possible use of the system.

Wessling, Francis C., III↗

NASA NTRS2023

Backup Optical Navigation Attitude for Artemis-1 Backup Attitude Ground Tool

The Backup Optical Navigation Attitude (BONA) software was developed as a response to the Artemis1 Power Distribution Unit (PDU) hardware problems found early in 2021. The hardware problem, a capacitor installed incorrectly, removes the intended redundant path for which the PDU can relay its power and data to attached devices. Specific to the BONA context, the concern is that a failure of the single remaining path in this PDU would lead to an inability to communicate with one of the two-star trackers (STs) on Orion. By flight rule, being reduced to a single star tracker means an immediate turn around end-of-mission. The BONA software is not intended to extend the mission, but instead act as a single star tracker attitude confirmation tool. Insurance, if you will, for the Orion project that a ground tool is available to compare the single remaining ST results with optical navigation images retrieved during the mission. Engineers operating BONA in the Mission Control Center (MCC) Mission Evaluation Room (MER) will analyze the downlinked star field images and based on the stars identified and known time of image, derive vehicle attitude estimates, which can be compared with the ST results. Potentially, in extreme situations, and with expert recommendation, the derivations could lead to navigation state updates being commanded to Orion.

GNC↗

NASA NTRS1986

Space infrared telescope pointing control system. Infrared telescope tracking in the presence of target motion

The use of charge-coupled-devices, or CCD's, has been documented by a number of sources as an effective means of providing a measurement of spacecraft attitude with respect to the stars. A method exists of defocussing and interpolation of the resulting shape of a star image over a small subsection of a large CCD array. This yields an increase in the accuracy of the device by better than an order of magnitude over the case when the star image is focussed upon a single CCD pixel. This research examines the effect that image motion has upon the overall precision of this star sensor when applied to an orbiting infrared observatory. While CCD's collect energy within the visible spectrum of light, the targets of scientific interest may well have no appreciable visible emissions. Image motion has the effect of smearing the image of the star in the direction of motion during a particular sampling interval. The presence of image motion is incorporated into a Kalman filter for the system, and it is shown that the addition of a gyro command term is adequate to compensate for the effect of image motion in the measurement. The updated gyro model is included in this analysis, but has natural frequencies faster than the projected star tracker sample rate for dim stars. The system state equations are reduced by modelling gyro drift as a white noise process. There exists a tradeoff in selected star tracker sample time between the CCD, which has improved noise characteristics as sample time increases, and the gyro, which will potentially drift further between long attitude updates. A sample time which minimizes pointing estimation error exists for the random drift gyro model as well as for a random walk gyro model.

Powell, J. D.↗

NASA NTRS2013

Three-Axis Attitude Estimation With a High-Bandwidth Angular Rate Sensor

A continuing challenge for modern instrument pointing control systems is to meet the increasingly stringent pointing performance requirements imposed by emerging advanced scientific, defense, and civilian payloads. Instruments such as adaptive optics telescopes, space interferometers, and optical communications make unprecedented demands on precision pointing capabilities. A cost-effective method was developed for increasing the pointing performance for this class of NASA applications. The solution was to develop an attitude estimator that fuses star tracker and gyro measurements with a high-bandwidth angular rotation sensor (ARS). An ARS is a rate sensor whose bandwidth extends well beyond that of the gyro, typically up to 1,000 Hz or higher. The most promising ARS sensor technology is based on a magnetohydrodynamic concept, and has recently become available commercially. The key idea is that the sensor fusion of the star tracker, gyro, and ARS provides a high-bandwidth attitude estimate suitable for supporting pointing control with a fast-steering mirror or other type of tip/tilt correction for increased performance. The ARS is relatively inexpensive and can be bolted directly next to the gyro and star tracker on the spacecraft bus. The high-bandwidth attitude estimator fuses an ARS sensor with a standard three-axis suite comprised of a gyro and star tracker. The estimation architecture is based on a dual-complementary filter (DCF) structure. The DCF takes a frequency- weighted combination of the sensors such that each sensor is most heavily weighted in a frequency region where it has the lowest noise. An important property of the DCF is that it avoids the need to model disturbance torques in the filter mechanization. This is important because the disturbance torques are generally not known in applications. This property represents an advantage over the prior art because it overcomes a weakness of the Kalman filter that arises when fusing more than one rate measurement. An additional advantage over prior art is that, computationally, the DCF requires significantly fewer real-time calculations than a Kalman filter formulation. There are essentially two reasons for this: the DCF state is not augmented with angular rate, and measurement updates occur at the slower gyro rate instead of the faster ARS sampling rate. Finally, the DCF has a simple and compelling architecture. The DCF is exactly equivalent to flying two identical attitude observers, one at low rate and one at high rate. These attitude observers are exactly of the form currently flown on typical three-axis spacecraft.

Bayard, David S.↗

NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Guidance, Navigation and Control (GN&C) Overview

The National Aeronautics and Space Administration s (NASA) Lunar Reconnaissance Orbiter (LRO) launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle and into a direct insertion trajectory to the oon. LRO, which was designed, built, and operated by the NASA Goddard Space Flight Center in Greenbelt, MD, is gathering crucial data on the lunar environment that will help astronauts prepare for long-duration lunar expeditions. The mission has a nominal life of 1 year as its seven instruments find safe landing sites, locate potential resources, characterize the radiation environment, and test new technology. To date, LRO has been operating well within the bounds of its requirements and has been collecting excellent science data images taken from the LRO Camera Narrow Angle Camera of the Apollo landing sites appeared on cable news networks. A significant amount of information on LRO s science instruments is provided at the LRO mission webpage. LRO s Guidance, Navigation and Control (GN&C) subsystem is made up of an onboard attitude control system (ACS) and a hardware suite of sensors and actuators. The LRO onboard ACS is a collection of algorithms based on high level and derived requirements, and reflect the science and operational events throughout the mission lifetime. The primary control mode is the Observing mode, which maintains the lunar pointing orientation and any offset pointing from this baseline. It is within this mode that all science instrument calibrations, slews and science data is collected. Because of a high accuracy requirement for knowledge and pointing, the Observing mode makes use of star tracker (ST) measurement data to determine an instantaneous attitude pointing. But even the star trackers alone do not meet the tight requirements, so a six-state Kalman Filter is employed to improve the noisy measurement data. The Observing mode obtains its rate information from an inertial reference unit (IRU) and in the event of an IRU failure, the rate data is be derived from the star tracker, but with degraded pointing performance. The Delta-V control mode responsibility is to maintain attitude pointing during the cruise trajectory, insertion burns and lunar orbit maintenance by adjustments made to the spacecraft s velocity magnitude and vector direction. The ACS also provides for a thruster based system momentum management algorithm (known as Delta-H) to maintain the system and wheel momentum to within acceptable levels. In the event an anomaly causes the LRO spacecraft to lose the ability to maintain its current attitude pointing, a Sun Safe mode is included in the ACS for the purpose of providing a known power and thermally safe coarse inertial sun attitude for an indefinite period of time, within the manageable limits of the reaction wheels. The Sun Safe mode is also the initial spacecraft control mode off of the launch vehicle and provides for a means to null tip-off rates immediately after separation. The nominal configuration is to use the IRU for rate information in the controller. In the event of a gyro failure a gyroless control mode was developed that computes rate information from the CSS data.

Garrick, Joseph↗