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Two state-variable representations derived for continuous-time plant driven by control algorithm including zero-order hold and measurements sampled at mutliple rates by multiple-input/multiple-output moving-average processes. New representations enhance observability and controllability of plant. Applications include mathematical modeling of navigation systems including star trackers, gyroscopes, and accelerometers.
Compared here are two methods of on-orbit alignment of vector attitude sensors. The first method uses the angular difference between simultaneous measurements from two or more sensors. These angles are compared to the angular differences between the respective reference positions of the sensed objects. The alignments of the sensors are adjusted to minimize the difference between the two sets of angles. In the second method, the sensor alignment is part of a state vector that includes the attitude. The alignments are adjusted along with the attitude to minimize all observation residuals. It is shown that the latter method can result in much less alignment uncertainty when gyroscopes are used for attitude propagation during the alignment estimation. The additional information for this increased accuracy comes from knowledge of relative attitude obtained from the spacecraft gyroscopes. The theoretical calculations of this difference in accuracy are presented. Also presented are numerical estimates of the alignment uncertainties of the fixed-head star trackers on the Extreme Ultraviolet Explorer spacecraft using both methods.
An earlier heuristic study of the fine attitude sensors for the Solar Maximum Mission (SMM) revealed a temperature dependence of the alignment about the yaw axis of the pair of fixed-head star trackers relative to the fine pointing Sun sensor. Here, new sensor alignment algorithms which better quantify the dependence of the alignments on the temperature are developed and applied to the SMM data. Comparison with the results from the previous study reveals the limitations of the heuristic approach. In addition, some of the basic assumptions made in the prelaunch analysis of the alignments of the SMM are examined. The results of this work have important consequences for future missions with stringent attitude requirements and where misalignment variations due to variations in the temperature will be significant.
The mission of the Hubble Space Telescope (HST) is to explore the expanding and evolving universe. During the 3,000 operating hours every year for the next 15 years or more, the HST will be used to study: galaxies; pulsars; globular clusters; neighboring stars where planets may be forming; binary star systems; condensing gas clouds and their chemical composition; and the rings of Saturn and the swirling ultraviolet clouds of Venus. The major technical achievements - its nearly perfect mirrors, its precise guidance system of rate gyroscopes, reaction wheels, star trackers, and fine guidance sensors are briefly discussed. The scientific instruments on board HST are briefly described. The integration of the equipment and instruments is outlined. The Space Telescope Science Institute (STScI) has approved time for 162 observations from among 556 proposals. The mission operation and data flow are explained.
The dislodgement, venting, and redeposition of particles on a surface caused by the vibroacoustic, gravitational, and aerodynamic forces are estimated and the resulting areal obscurations of the surface are calculated. The data on particle redistribution estimated to occur during a Shuttle launch are employed to estimate the obscuration of a spacecraft's star tracker, as it is carried into orbit by the Shuttle. The approach used for that calculation is generalized so that it can be employed for other applications where either the Shuttle or an Expendable Launch Vehicle (ELV) is employed for launch. Approaches for evaluating particle redistribution for other applications and for general use are indicated.
A detailed review of the in-flight calibration of the Hubble Space Telescope attitude sensors is given. The review, which covers the period from the April 24, 1990 launch of the spacecraft until April 1991, describes the calibrations required and accuracies achieved for the four principal attitude sensing systems on the spacecraft: the magnetometers, the fixed-head star trackers, the gyroscopes, and the fine guidance sensors.
This paper presents a comprehensive approach to filter and dynamics modeling for attitude determination error analysis. The discussion includes models for both batch least-squares and sequential estimators, a specific dynamic model for attitude determination error analysis of a three-axis stabilized spacecraft equipped with strapdown gyros, and the incorporation of general attitude sensor observations. An analyst using this approach to perform an error analysis chooses a subset of the spacecraft parameters to be 'solve-for' parameters, which are to be estimated, and another subset to be 'consider' parameters, which are assumed to have errors but not to be estimated. The result of the error analysis is an indication of overall uncertainties in the 'solve-for' parameters, as well as the contributions of the various error sources to these uncertainties, including those of errors in the a priori 'solve-for' estimates, of measurement noise, of dynamic noise (also known as process noise or plant noise), and of 'consider' parameter uncertainties. The analysis of attitude, star tracker alignment, and gyro bias uncertainties for the Gamma Ray Observatory spacecraft provide a specific example of the use of a general-purpose software package incorporating these models.
At about 14,800 km from the Comet Halley nucleus, on the inbound leg, at least six of the sensors onboard the Giotto spacecraft observed an unusual, brief (about 30 to 500 ms) event: the ion-mass spectrometer data show a brief flow of energetic (up to several hundred electron volts) plasma consisting of protons, water group, and heavier ions. The Johnstone plasma analyzer data show a short burst of plasma, while the dust impact detector system data show an impact event in four of its detectors. The magnetometer signature of the event shows two brief dips in the field. The sudden change in the spacecraft attitude and spin rate observed by the camera at that same time has been interpreted as the result of a large (5 mg or more) dust-particle impact on the front bumper shield of the spacecraft. In addition, at about the same time the spacecraft star-tracker suffered damage. The report combines direct measurements of the composition and dynamics of a dust-impact plasma cloud, the dust particle mass, and the location of the impact on the spacecraft. Analysis of the data indicate that the impacting particle was water or ice-bearing, possibly loosely compared, and was composed of one or more of: carbon, nitrogen, and silicon.
The attitude determination sensing and processing of the system are considered, and inertial reference units, star trackers, and beacons and landmarks are discussed as well as an extended Kalman filter and expected attitude-determination performance. The baseline controller is overviewed, and a spacecraft motion compensation (SMC) algorithm, disturbance environment, and SMC performance expectations are covered. Detailed simulation results are presented, and emphasis is placed on dynamic models, attitude estimation and control, and SMC disturbance accommmodation. It is shown that the attitude control system employing gyro/star tracker sensing and active three-axis control with reaction wheels is capable of maintaining attitude errors of 1.7 microrad or less on all axes in the absence of attitude disturbances, and that the sensor line-of-sight pointing errors can be reduced to 0.1 microrad by SMC.
A scheme for estimating the angular velocity of a spacecraft without rate gyros is presented. It is based upon a nonlinear estimator whose inputs are measured inertial vectors and their calculated time derivatives relative to vehicle axes. It works for all spacecraft attitudes and requires no knowledge of attitude. It can use measurements from a variety of onboard sensors like Sun sensors, star trackers, or magnetometers, and in concert. It can also use look angle measurements from onboard tracking antennas for tracking and data relay satellites or global positioning system satellites. In this paper, it is applied to a Sun point scheme on the Hubble Space Telescope assuming all or most of its onboard rate gyros have failed. Simulation results are presented for verification.
The results of on-orbit calibration for several satellites by the flight Dynamics Facility (FDF) at GSFC are reviewed. The examples discussed include attitude calibrations for sensors, including fixed-head star trackers, fine sun sensors, three-axis magnetometers, and inertial reference units taken from recent experience with the Compton Gamma Ray observatory, the Upper Atmosphere Research Satellite, and the Extreme Ultraviolet Explorer calibration. The methods used and the results of calibration are discussed, as are the improvements attained from in-flight calibration.
The results of a student project to design a lunar lander in the context of a specifically defined mission are presented. The Lunar Scout will be launched from Cape Canaveral, Florida onboard a Delta II launch vehicle. The Delta II will carry the lander and its payload to a 1367 km orbit. Once it reaches that altitude, a STAR 48A solid rocket motor will kick the spacecraft into a lunar trajectory. After burnout of the lunar insertion motor, it will be jettisoned from the spacecraft. The flight from the earth to the moon will take approximately 106.4 hours. During this time the battery, which was fully charged prior to launch, will provide all power to the spacecraft. Every hour, the spacecraft will use its sun sensors and star trackers to update its position, maintain some stabilization and relay it back to earth using the dipole antennas. At the start of its lunar trajectory, the spacecraft will fire one of its 1.5 N thrusters to spin in at a very small rate. The main reason for this is to prevent one side of the spacecraft from overheating in the sun. When the spacecraft nears the moon, it will orient itself for the main retro burn. At an altitude of 200 km, a 4400 N bipropellant liquid thruster will ignite to slow the spacecraft. During the burn, the radar altimeter will be turned on to guide the spacecraft. The main retro rocket will slow the lander to 10 m/s at an approximate altitude of 40 km above the moon. From there, the space craft will use four 4.5 N hydrazine vertical thrusters and 1.5 N horizontal thrusters to guide the spacecraft to a soft landing. Once on the ground, the lander will shutoff the radar and attitude control systems. After the debris from the impact has settled, the six solar panels will be deployed to begin recharging the batteries and to power up the payload. The feedhorn antenna will then rotate to fix itself on the earth. Once it moves, it will stay in that position for the spacecraft's lifetime. The payload will then be activated to begin the lunar mission.
This paper will discuss in detail the relative navigation system requirements and sensor trade-offs for Automatic Rendezvous and Capture. Rendezvous navigation filter development will be discussed in the context of navigation performance requirements for a 'Phase One' AR&C system capability. Navigation system architectures and the resulting relative navigation performance for both cooperative and uncooperative target vehicles will be assessed. Relative navigation performance using rendezvous radar, star tracker, radiometric, laser and GPS navigation sensors during appropriate phases of the trajectory will be presented. The effect of relative navigation performance on the Integrated AR&C system performance will be addressed. Linear covariance and deterministic simulation results will be used. Evaluation of relative navigation and IGN&C system performance for several representative relative approach profiles will be presented in order to demonstrate the full range of system capabilities. A summary of the sensor requirements and recommendations for AR&C system capabilities for several programs requiring AR&C will be presented.
The Arthur Holly Compton Gamma Ray Observatory (GRO) was launched by the shuttle Atlantis in April 1991. This paper presents the results of the attitude sensor calibration that was performed during the early mission. The GSFC Flight Dynamics Facility (FDF) performed an alignment calibration of the two fixed-head star trackers (FHST's) and two fine Sun sensors (FSS's) on board Compton GRO. The results show a 27-arcsecond shift between the bore sights of the FHST's with respect to prelaunch measurements. The alignments of the two FSS's shifted by 0.20 and 0.05 degree. During the same time period, the Compton GRO science teams performed an alignment calibration of the science instruments with respect to the attitude reported by the on board computer (OBC). In order to preserve these science alignments, FDF adjusted the overall alignments of the FHST's and FSS's, obtained by the FDF calibration, such that when up linked to the OBC, the shift in the OBC-determined attitude is minimized. FDF also calibrated the inertial reference unit (IRU), which consists of three dual-axis gyroscopes. The observed gyro bias matched the bias that was solved for by the OBC. This bias drifted during the first 6 days after release. The results of the FDF calibration of scale factor and alignment shifts showed changes that were of the same order as their uncertainties.
The Upper Atmosphere Research Satellite (UARS) uses fixed-head star trackers (FHST's) and inertial reference units (IRU's) to determine and control its attitude. This combination of fine sensors results in attitude knowledge accuracies to better than 10 arc-seconds (1 sigma). UARS also has a variety of coarse attitude sensors on board: the three-axis magnetometer (TAM), the coarse Sun sensor (CSS), the fine Sun sensor (FSS), and the Earth sensor assembly (ESA). By comparing attitude solutions using coarse sensors with FHST-determined attitude solutions, estimates can be made of the accuracy of the coarse sensors. This paper presents the results of an analysis that compares attitude solutions using various combinations of UARS coarse attitude sensor data with FHST attitude solutions.
The design of the DSPSE Satellite Controller (DSC) is baselined as a single-string satellite controller. The DSC performs two main functions: health and maintenance of the spacecraft; and image capture, storage, and playback. The DSC contains two processors: a radiation-hardened Mil-Std-1750, and a commercial R3000. The Mil-Std-1750 processor performs all housekeeping operations, while the R3000 is mainly used to perform the image processing functions associated with the navigation functions, as well as performing various experiments. The DSC also contains a data handling unit (DHU) used to interface to various spacecraft imaging sensors and to capture, compress, and store selected images onto the solid-state data recorder. The development of the DSC evolved from several key requirements; the DSPSE satellite was to do the following: (1) have a radiation-hardened spacecraft control system and be immune to single-event upsets (SEU's); (2) use an R3000-based processor to run the star tracker software that was developed by SDIO (due to schedule and cost constraints, there was no time to port the software to a radiation-hardened processor); and (3) fly a commercial processor to verify its suitability for use in a space environment. In order to enhance the DSC reliability, the system was designed with multiple processing paths. These multiple processing paths provide for greater tolerance to various component failures. The DSC was designed so that all housekeeping processing functions are performed by either the Mil-Std-1750 processor or the R3000 processor. The image capture and storage is performed either by the DHU or the R3000 processor.
The promise and technical feasibility of a GPS sensor updating inertial attitude for the guidance, navigation, and control (GN&C) and the flight control system functions of spacecraft are demonstrated. Attitude control simulations are reported which suggest that GPS/INS systems will be as suitable as astroinertial GN&C systems of future space vehicle navigation. However, arcsec attitude accuracy star trackers will still be required for special payloads. HT need for international cooperation and agreements before worldwide implementation can be achieved is emphasized.
A rugged, miniaturized, optical cross-correlator that recognizes a single object is particularly suitable for performing a single-vision function, such as pattern recognition for semi-autonomous navigation, landing, and docking of vehicles to a pre-designated landing mark. The optical cross-correlator, with a video input from a simple imaging system and the output of the optical correlation plane processed using the standard star tracker software, produces sufficient information for a spacecraft's terminal homing navigation system to complete a docking maneuver.