Hughes Danbury Optical Systems, Inc. (HDOS) is developing a new generation of star trackers that combine low-weight and power consumption with high performance, high reliability, and survivability. Our Miniature Star Tracker (MST), Model HD-1003, represents our next step in the evolution of our advanced star tracker (ASTRA) family of CCD-based star trackers and stresses compact packaging to minimize the impact on the spacecraft design. Small size lowers tracker procurement, integration and launch costs, and provides the user with a high degree of flexibility in selecting the optimum attitude sensor suite, especially for small-sats. Weighing less than 6 lbs, our MST incorporates the latest detector, processor, and application- specific integrated circuit (ASIC) electronics technology to provide exceptional accuracy and survivability in the natural, proton-rich space environment.
Hughes Danbury Optical Systems (HDOS) has built and successfully tested an Engineering Model (EM) Star Tracker for the Attitude Determination System (ADS) of the Space Station Freedom. This work was accomplished under a contract from Honeywell's Space Systems Group at Clearwater, Florida. The EM has empirically demonstrated that the design is consistent with functional and performance features of the flight model, including the ability to acquire and track up to four stars in the debris rich environment surrounding the Space Station. Included is a discussion of the modular design of the tracker which lends itself to thorough subassembly testing. Finally the integrated testing and test results are presented.
The HD-1002 (also known as MACS) solid state star trackers being built by Hughes Danbury Optical Systems, Inc. (HDOS) for Martin Marietta Astro Space Division for use in their Modular Attitude Control Systems (MACS) Module are improved and modified versions of the ASTRA1 star trackers now in use on board the TOPEX/POSEIDON satellite. The ASTRA1 design was based on the pioneering work accomplished at HDOS over the past decade. Along with the set of trackers being built by HDOS for Space Station Freedom, these trackers answer a variety of application requirements for spacecraft attitude control systems. This paper addresses the main features of the MACS trackers, their role in the MACS Module, and summarizes the excellent preliminary performance results of the tracker, as supported by measured test data.
The ASTROS Star Tracker (AST), designed and built by JPL as a demonstration instrument in the early 80s was integrated with the astro scientific payload in 1985 for a planned March 1986 launch. After a slip of nearly five years, it was finally launched in December 1990, gathering extensive scientific data for over one hundred scientific targets. This paper reviews some of the AST results from the Astro-1 flight. After the five-year stand down, photometric, spectral, and geometric calibrations remained virtually unchanged, allowing predictable performance on all targets and successful automatic identification of every star field. Although small changes in the optical point-spread function increased the centroid error, this did not affect operation for Astro, and should be correctable for future instruments. Our data suggest that calibration of centroid error to the 1/100 pixel level is achievable when the point-spread function remains stable. The data are also consistent with the noise-equivalent-angle (NEA) of 1/300 pixel measured in the laboratory for bright stars.
The TOPEX/POSEIDON spacecraft was launched on August 10, 1992. This paper presents data on the measured performance of the ASTRA Star Trackers supplied by Hughes Danbury Optical Systems (HDOS) for this satellite. The HDOS ASTRA Star Tracker is a charge coupled device (CCD), microprocessor based replacement for the NASA Standard Fixed Head Star Tracker. The position and magnitude accuracy of the star trackers computed from measured flight data are compared with ground measurements and system models. The performance of novel transient rejection algorithms implemented in the ASTRA Star Tracker which allows uninterrupted operation in the South Atlantic Anomaly (SAA) where the sensor is subjected to high proton flux levels, also are presented.
The Space Station Attitude Determination system (ADS) will utilize star trackers with CCD optical sensors and inertial sensor assemblies (ISA) with ring laser gyro (RLG) inertial rotation sensors. The ADS will manipulate star data from the star trackers--star position and magnitude--to autonomously initialize the space station attitude and to provide periodic updates to a Kalman filter during normal operation. This paper describes the functional interfacing between the star trackers and the Attitude Determination Function (ADF)--ADS central processor resident software--that provides for the exchange of data and commands between the star tracker and the ADF. This paper also describes the process whereby the data received from the star tracker is processed for generation of an initial attitude for delivery to the Kalman filter for an attitude update.
A ground-based software system to calibrate the attitude control sensors for the Ocean Topography Experiment (TOPEX) spacecraft is described. The algorithm determines sensor misalignment, bias and scale factor errors from gyro, sun sensor and star tracker measurements. The inherent yaw slew motion of the spacecraft during normal mission mode is exploited to make the error parameters observable. A two loop recursive least-squares algorithm is implemented with the feature of inhibiting the update of the error parameter state until all of the available data has been processed. This feature eliminates the estimation state feedback typical of other algorithms which can cause instability and convergence problems.
Although the analytical groundwork for understanding two-dimensional object images and various aspects of computer vision has been laid, we have not yet applied these concepts to automating the process of obtaining science images during space exploration missions. Our current approach in specifying pointing-command sequences relies heavily on target predicts, based on predicted target and spacecraft ephemerides, that propagate the target position as a function of time during pointing operations. However, because the actual position of the target is never measured on board, the tracking loop must be closed through the ground processing operation. This round-trip communication-time limit can place severe limits on pointing accuracy, particularly during brief moments of closest approach. In this paper, we formulate and outline autonomous feature-based pointing requirements that can support a wide range of space science missions. The issues addressed include target-position estimation; acquisition and tracking of unresolved, partially viewed, and close-up targets; operation, and validation of point designs. Detailed algorithms and test results based on Voyager image data are also presented.
Spatial acquisition and precision beam pointing functions are critical to spaceborne laser communication systems. A recent system study indicated that a single high bandwidth CCD detector can be used to perform both spatial acquisition and tracking functions. Compared to previous lasercom hardware design, the array tracking concept offers reduced system complexity by reducing the number of optical elements in the design. Specifically, the design requires only one detector and one beam steering mechanism. It also provides means to optically close the point-ahead control loop. The technology required for high bandwidth array tracking was examined and shown to be consistent with current state of the art. It is believed that the single detector design can lead to a significantly reduced system complexity and a lower system cost.
In the past, NASA tracking systems which provided range and bearing to targets have primarily been radar based. Advanced projects such as unmanned missions to the moon and Mars need automated rendezvous and capture (AR&C) to reduce operating costs and improve mission reliability. In the Tracking Techniques Branch at Johnson Space Center we are investigating the feasibility of a stereo digital image based sensor for AR&C. This system differs from traditional stereo systems in two significant ways. First, a passive cooperative target is used to calculate the range to three specific points as opposed to calculating a range map of the entire scene. The second unique feature is the image processing algorithms that are used to identify the target. In this paper the sensor's performance is compared to preliminary AR&C operating range and accuracy requirements. A theoretical error model is presented that predicts the sensor's accuracy as a function of range.
This paper summarizes microgravity experiments for space platforms and presents test results of a three-axes active isolation demonstration system that has been built and tested. This mount has demonstrated isolation capability down to 0.05 Hz nd a transmissibility curve which provides isolation levels necessary for microgravity and precision pointing aboard space platforms. The system has high-bandwidth inertial servo loops using accelerometers for active disturbance rejection to isolate the payload from external torque disturbances by the utility transfer devices and moving mechanisms.
Many of the experiments to be performed within the various space agencies' microgravity programs are extremely sensitive to low frequency spacecraft vibration. The microgravity isolation mount (MGIM) consists of a free floating platform, accommodated within the spacecraft's experiment racks, which isolates sensitive payloads from ambient disturbances. This paper describes the main features of the MGIM and discusses some of the factors considered when deciding on a control system strategy. The emphasis of the paper is on the trade-off which occurs between the heat dissipation capacity of the thermal subsystem and the achievable microgravity level. Thermal dissipation from the platform is by means of interleaved cooling fins located on both the platform and its enclosure. Theoretical expressions are developed for the force caused by air motion in the gap between the fins and experimental results which verify these expressions are presented. It is shown, by means of computer simulation, that the attainable microgravity quality is limited by the dimensions of the cooling fins required to support a specific payload heat dissipation capacity.
Spacecraft capable of carrying modest to intermediate size science payloads into Earth orbit at relatively low cost are being investigated by the Marshall Space Flight Center at the request of the Astrophysics and Space Physics Division of OSSA. Intermediate-class space science missions, such as the Lunar Ultraviolet Transit Experiment (LUTE), Inner Magnetosphere Imager (IMI), the Solar Ultraviolet Radiation and Correlative Emissions (SOURCE) experiment, and the Long Duration Exposure Facility (LDEF-II) are expected to have a progressively larger role in NASA's space science program into the next century. These and other space science missions have been examined to define the systems, subsystems, and interface requirements needed to accomplish their stated objectives. This paper discusses the science objectives, technical requirements and major issues posed by IMI, LUTE, SOURCE, and LDEF-II and will address MSFC's new ways of doing business.
A conceptual attitude control subsystem design for the Pluto Fast Flyby spacecraft is described. Mass, cost, schedule and performance, approximately in that order, drove the mission, spacecraft, as well as the attitude control subsystem design. The paper discusses the key mission requirements impacting the attitude control subsystem design, as well as the important subsystem trades. The spacecraft is a three axis stabilized vehicle using cold gas jets for attitude control and hydrazine thrusters for trajectory correction maneuvers. Attitude determination relies heavily on a low mass star tracker capable of determining attitude by pointing anywhere in the celestial sphere. Tracking of planetary features with the star tracker may also be desirable. A small inertial reference unit and a sun sensor will accompany the tracker to complete the suite of components for attitude determination.
This paper describes the design of a stabilization system for control of a spacecraft designed to be launched on a medium-sized launch vehicle. The system is required to provide accurate attitude control and determination for a nominally nadir-pointed spacecraft with moderate attitude maneuver capability. The spacecraft uses reaction wheels as the primary actuators and an effective magnetic momentum desaturation system. Attitude determination is based on stellar position measurements and gyroscope bias estimations to maintain attitude knowledge. A significant design problem involves the stabilization of maneuverable flexible solar arrays with low natural frequency. The attitude determination and control system is composed of sensors, actuators and an integrated electronics unit. A propulsion system allows for incremental velocity adjust capability and actuation during back-up attitude control. This paper provides a description of the attitude determination and control system configuration and functional operation. In addition the component characteristics are presented.
An automatic star identification program, which has previously been demonstrated only in computer simulation, has been integrated and tested with the tracker test lab prototype star tracker at Table Mountain Observatory. Thirteen of sixteen sky frames were matched correctly with no a priori attitude knowledge. The cause of failure in the frames that failed was determined to be algorithmic, and the flaws in the algorithm causing the failure were identified. The methodology of both Table Mountain Observatory test runs and computer simulated test runs are described, as well as results and recommendations.
Inertially referenced attitude determination for pointing and control is a well established need which demands continual optimization. Star trackers can accomplish the task by direct stellar measurements. This paper presents the results of design trades for Hughes Danbury Optical System's (HDOS) Advanced Star Tracker (ASTRA) products with the built in capability to output inertial attitude, typically in the form of a quaternion in an earth centered inertial (ECI) coordinate system. This is an enhanced capability from standard tracker products which only output star image centroid positions. Attitude accuracy predictions for a variety of trades have been evaluated using an orbital flight simulator. This simulator integrates star tracker performance modeling with algorithms for attitude determination. Trades include field of view (FOV) size, sensitivity, number of stars tracked and star catalog size.
The Attitude Determination System (ADS) aboard Space Station Freedom will utilize star trackers (ST) with CCD optical sensors and inertial sensor assemblies (ISA) with ring laser gyro (RLG) inertial rotation sensors. An on-board star catalog of some 2100 stars, with magnitudes extending up to 5.15 instrument magnitude, will be used in conjunction with the ADS. The catalog contains both stellar inertial position and instrument magnitude data. This paper describes the procedure used to estimate the stellar instrument magnitudes for stars in the catalog, and characterizes the associated errors. Basically, the procedure involves the use of published detailed stellar spectral irradiance data and ST vendor provided optical and CCD spectral responsivity curves to numerically determine the instrument magnitudes of the stars associated with the published data.
NASA has an ambitious plan for exploring the universe involving
many space-based observatories. The CSI program (Control
Structure Interaction) is a NASA funded project to design
microprecision spacecraft observatories from a multidisciplinary
approach by integrating the structural, control and optical
subsystems of these spacecraft. A Focus Mission lnterferometer
(FMI), representative of one of NASA's astrometric missions for
extrasolar planet detection was selected as the focus for developing
this multidisciplinary technology. Very early in this study it was
discovered that integrated modeling tools were necessary to predict
the on orbit performance of the FMI.
A new radar glint model has been created to evaluate the terminal characteristics of the Advanced Medium Range Air-to-Air Missile (AMRAAM). This model was designed to use deterministically based glint parameters that develop accurate glint characteristics for real-time simulations. The model uses a deterministically derived data table to specify the first order glint statistics; it then generates random glint characteristics which are a function of target type, target orientation, and the target rotation rate. This combination allows the model to develop radar cross section and glint data to drive the real-time AMRAAM simulation program.
Charge-coupled devices (CCDs) have been used extensively in the past in star trackers and fine guidance systems. A new technology, the active pixel sensor, is a possible successor to CCDs. This technology potentially features the same sensitivity and performance of the CCD with additional improvements. These improvements include random access capability, easy window-of- interest readout, non-destructive readout for signal-to-noise improvement, high radiation tolerance, simplified clocking voltages, and easy integration with other on-chip signal processing circuitry. The state-of-the-art of this emerging technology and its potential application to guidance and navigation systems is discussed.
INSAT-2 series satellites are second generation, multipurpose satellites developed by Indian Space Research Organisation for telecommunication and meteorological remote sensing. INSAT-2A was launched successfully in July 1992. The solar sail mounted on the north face of the satellite balances the differential solar radiations torque on the solar panel mounted on the south face. The sail, being lightweight, during its deployment causes very little attitude disturbance on the satellite to sense the deployment. A novel utilization of an earth sensor (ES) mounted on the same face of the satellite gave a clear indication of the deployment initiation and the progress of the boom extension during this period. This paper briefly describes the observations made using the earth sensor during the sail deployment and analyzes the data received in terms of the sail extension in this period.
The Space Infrared Telescope Facility mission provides exciting pointing and control challenges. This paper describes the pointing and control subsystem (PCS) designed to meet these challenges. Cost, mass, performance, and life-time drove the choice of the new solar orbit and the re-design of the telescope, spacecraft, and the pointing and control subsystem. The PCS performs a number of key functions for the observatory, including pointing to and holding inertial targets, tracking moving targets, performing prescribed motion patterns, maintaining pointing constraints and providing safe hold during failure. To attain the high precision demanded by the observatory, the PCS relies on a sub-arcsecond star tracker and a 0.001 arcsecond resolution inertial reference unit. An affordable PCS is realized by locating the fine guidance sensor externally, utilizing a simple fault protection strategy, and maximizing the commonality in the use of hardware and in the development of flight software.
The following paper concerns the study of the Attitude Control System (ACS) of the Orbital Space Interferometer (OSI) proposed by Jet Propulsion Laboratory (JPL) of Pasadena, both for NASA's Astrometric Interferometry Mission (AIM) and for phase one of NASA's Towards Other Planetary Systems (TOPS) program.