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The METEOSAT ranging signal's available every 3 hours for about 1.5 minutes, its phase is measured against the second pulses of the station clocks. A system study shows that the accuracy obtainable could be in the order of 30 ns. The receiver unit consists of a 2 m diameter fixed antenna, a low noise broadband receiver, and a phase tracker and demodulator. The overall costs will be less than $20.000 US.
A sensitive sounding rocket instrument for moderate (about 10-A) resolution far-UV (1160-1750-A) spectrophotometry of faint astronomical objects has been developed. The instrument employes a photon-counting microchannel plate imaging detector and a concave grating spectrograph behind a 40-cm Dall-Kirkham telescope. A unique remote-control pointing system, incorporating an SIT vidicon aspect camera, two star trackers, and a tone-encoded command telemetry link, permits the telescope to be oriented to within 5 arc sec of any target for which suitable guide stars can be found. The design, construction, calibration, and flight performance of the instrument are discussed.
The Shuttle Infrared Telescope Facility (SIRTF) is being designed as a 1-m, cryogenically cooled telescope capable of a thirty-fold improvement over currently available infrared instruments. The SIRTF, mounted in the Orbiter bay on the Instrument Pointing System (IPS), requires that the image at the focal plane be stabilized to better than 0.1 arcsec with an absolute accuracy of 1 arcsec in order to attain this goal. Current estimates of IPS performance for both stability and accuracy indicate that additional stabilization will be necessary to meet the SIRTF requirements. An Image Motion Compensation (IMC) system, utilizing a Charge Coupled Device (CCD) star tracker located at the focal plane and a steerable mirror in the SIRTF optical path, has been designed to work in conjunction with the IPS.
A unique approach to onboard processing was developed that is capable of acquiring high quality image data for users in near real time. The approach is divided into two steps: the development of an onboard cloud detection system; and the development of a landmark tracker. The results of these two developments are outlined and the requirements of an operational guidance and control system capable of providing continuous estimation of the sensor boresight position are summarized.
The SMM attitude ground support system (AGSS) supports the acquisition of spacecraft roll attitude reference, performs the in-flight calibration of the attitude sensor complement, supports onboard control autonomy via onboard computer data base updates, and monitors onboard computer (OBC) performance. Initial roll attitude acquisition is accomplished by obtaining a coarse 3 axis attitude estimate from magnetometer and Sun sensor data and subsequently refining it by processing data from the fixed head star trackers. In-flight calibration of the attitude sensor complement is achieved by processing data from a series of slew maneuvers designed to maximize the observability and accuracy of the appropriate alignments and biases. To ensure autonomy of spacecraft operation, the AGSS selects guide stars and computes sensor occultation information for uplink to the OBC. The onboard attitude control performance is monitored on the ground through periodic attitude determination and processing of OBC data in downlink telemetry. In general, the control performance has met mission requirements. However, software and hardware problems have resulted in sporadic attitude reference losses.
A computer-controlled automatic facility containing two interactive microprocessors, peripherals, and an optical bench was used to obtain device characteristics for use in evaluating CID's as the sensor element in a prospective star tracker application. Each of the ST-256 chips tested contains a 256 x 256 pixel element array with 65536 picture elements. Two ST-256 devices were tested at 20 C under light levels of 10%, 0.1%, and 0.01% saturation, and in darkness. The saturation level was established at 0.95 microwatts sq cm as measured by the radiometeric sensor. Two test runs under each condition were performed for each chip. Two complete sets of data for each of the devices were thus obtained. A separate data field was used to store the results of each test condition. Repetitive runs of a test condition were stored sequentially in the same file. Examination of the data using print-outs is shown in examples.
A solar tracker for a solar collector is described in detail. The collector is angularly oriented by a motor wherein the outputs of two side-by-side photodetectors are discriminated as to three ranges: a first corresponding to a low light or darkness condition; a second corresponding to light intensity lying in an intermediate range; and a third corresponding to light above an intermediate range, direct sunlight. The first output drives the motor to a selected maximum easterly angular position; the second enables the motor to be driven westerly at the Earth rotational rate; and the third output, the separate outputs of the two photodetectors, differentially controls the direction of rotation of the motor to effect actual tracking of the Sun.
The objective of this experiment was to demonstrate electrical power generation using a small free-piston Stirling engine and linear alternator in conjunction with a parabolic solar collector. A test bed collector, formerly used at the JPL Table Mountain Observatory, was renovated and used to obtain practical experience and to determine test receiver performance. The collector was mounted on a two-axis tracker, with a cold water calorimeter mounted on the collector to measure its efficiency, while a separate, independently tracking radiometer was used to measure solar insolation. The solar receiver was designed to absorb energy from the collector, then transfer the resulting thermal energy to the Stirling engine. Successful testing of receiver/collector assembly yielded valuable inputs for design of the Stirling engine heater head.
The possible use of an optical for high rate data transmission from a deep space vehicle to an Earth-orbiting relay satellite while RF links are envisioned for the relay to Earth link was studied. A preliminary link analysis is presented for initial sizing of optical components and power levels, in terms of achievable data rates and feasible range distances. Modulation formats are restricted to pulsed laser operation, involving bot coded and uncoded schemes. The advantage of an optical link over present RF deep space link capabilities is shown. The problems of acquisition, pointing and tracking with narrow optical beams are presented and discussed. Mathematical models of beam trackers are derived, aiding in the design of such systems for minimizing beam pointing errors. The expected orbital geometry between spacecraft and relay satellite, and its impact on beam pointing dynamics are discussed.
The method of torque equilibrium attitude control used to control the reentry of Skylab to an altitude below 150 km without the use of thruster fuel once the attitude was established is discussed. The Skylab attitude and pointing control system, which included rate gyros, sun sensors, star tracker, the Apollo telescope mount digital computer, control moment gyros and cold-gas attitude thrusters, is presented. The 12 torque equilibrium attitudes found at which aerodynamic, gravity gradient and gyroscopic torques would balance are indicated, and the three of those at which the solar power supply would be adequate for attitude control are illustrated. The equilibrium seeking method employed is then examined, and the operation and performance of the torque equilibrium attitude control system during the three weeks prior to Skylab reentry are discussed. It is concluded that the torque equilibrium attitude control method developed for Skylab was successful in performing its assigned mission, and will be valuable for the design of future, low-altitude spacecraft or tethered vehicles.
The atmospheric emissions photometric imaging experiment (AEPI) to be flown on Spacelab 1 is designed to study faint natural and artificial atmospheric emission phenomena. Optical emissions are imaged in the region 2150 A to 7320 A using a television system consisting of two optical channels, one wide-angle and one telephoto. The detection system is an image-enhanced SEC vidicon. A third optical channel images onto the photocathode of a microchannel plate photomultiplier tube that has 100 discrete anodes. Photons are counted for each discrete anode, providing a direct measure of the luminosity of an object viewed by the TV telephoto lens, albeit with low spatial resolution. The AEPI detector is mounted on a two-axis gimbal comprised of a Modified Apollo Telescope Mount Star Tracker (MAST), which provides experiment pointing over a 40-deg x 80-deg range, exclusive of restrictions due to the proximity of other experiments. The pointing stability is 1 arcmin with respect to the spacecraft coordinate system for an exposure of 1 second. The tracking capability is 3.5 deg/s with a stability of 1 arcmin. The detector and pointing system are located on the Spacelab pallet. The experiment is controlled by stored programs resident in the Dedicated Experiment Processor located in the Spacelab module.
Proposed target tracking and ranging system uses two automatic video target trackers to keep two TV cameras trained on object being tracked. Microcomputer calculates range and range-rate information by triangulation. Input data for calculation are position coordinates of two cameras and pan and tilt aiming angles of two cameras. System is useful for target ranging at distances up to about 1,000 feet (300 m) in such applications as vehicle collision avoidance, traffic monitoring and surveillance. Also substitutes for short-range radar in situations where radar signal can not be tolerated.
Twelve engineering type tests were performed on several subsystems and experiment(s) of the OAO 3 spacecraft near its end of mission. The systems tested include: Princeton experiment package (PEP), fine error system guidance, inertial reference unit, star trackers, heat pipes, thermal control coatings, command and data handling, solar array; batteries, and onboard processor/power boost regulator. Generally, the systems performed well for the 8 1/2 years life of OAO 3, although some degradation was noted in the sensitivity of PEP and in the absorptivity of the skin coatings. Battery life was prolonged during the life of the mission in large part by carefully monitoring the charge-discharge cycle with careful attention not to overcharge.
A baseline study for a generalized spacecraft attitude calibration/verification system is presented. It can be used to define software specifications for three major functions required by a mission: the pre-launch parameter observability and data collection strategy study; the in-flight sensor calibration; and the post-calibration attitude accuracy verification. Analytical considerations are given for both single-axis and three-axis spacecrafts. The three-axis attitudes considered include the inertial-pointing attitudes, the reference-pointing attitudes, and attitudes undergoing specific maneuvers. The attitude sensors and hardware considered include the Earth horizon sensors, the plane-field Sun sensors, the coarse and fine two-axis digital Sun sensors, the three-axis magnetometers, the fixed-head star trackers, and the inertial reference gyros.
The problem of designing lateral control systems for the safe landing of VTOL aircraft on small ships is addressed. A ship model is derived. The issues of estimation and prediction of ship motions are discussed, using optimal linear linear estimation techniques. The roll motion is the most important of the lateral motions, and it is found that it can be predicted for up to 10 seconds in perfect conditions. The automatic landing of the VTOL aircraft is considered, and a lateral controller, defined as a ship motion tracker, is designed, using optimal control techniqes. The tradeoffs between the tracking errors and the control authority are obtained. The important couplings between the lateral motions and controls are demonstrated, and it is shown that the adverse couplings between the sway and the roll motion at the landing pad are significant constraints in the tracking of the lateral ship motions. The robustness of the control system, including the optimal estimator, is studied, using the singular values analysis. Through a robustification procedure, a robust control system is obtained, and the usefulness of the singular values to define stability margins that take into account general types of unstructured modelling errors is demonstrated. The minimal destabilizing perturbations indicated by the singular values analysis are interpreted and related to the multivariable Nyquist diagrams.
A method is described for visual object tracking by a computer vision system using TV cameras and special low-level image processing hardware. The tracker maintains an internal model of the location, orientation, and velocity of the object in three-dimensional space. This model is used to predict where features of the object will lie on the two-dimensional images produced by stereo TV cameras. The differences in the locations of features in the two-dimensional images as predicted by the internal model and as actually seen create an error signal in the two-dimensional representation. This is multiplied by a generalized inverse Jacobian matrix to deduce the error in the internal model. The procedure repeats to update the internal model of the object's location, orientation and velocity continuously.
In August 1980, the Second High Energy Astronomy Observatory suffered the simultaneous double failure of two gyros, leaving the observatory with one gyro less than what is required for operation. The function of the missing gyro was replaced to allow restoration of spacecraft operation by reprogramming the on-board computer to derive rate from star and sun sensors, and by utilizing complex procedures. Large sun sensor quantization, jump characteristics of the star tracker, and extremely limited computer memory capacity further complicated the problem. Recovery attempts continued for over three months, during which time a suitable rate algorithm was derived. Final verification of the technique was never completed, however, due to the unexpected revival of one of the failed gyros.
The Voyager attitude and articulation control system is designed for a three-axis stabilized spacecraft; it uses a biasable sun sensor and a Canopus Star Tracker (CST) for celestial control, as well as a dry inertial reference unit, comprised of three dual-axis dry gryos, for inertial control. A series of complex maneuvers was required during the first of two Voyager spacecraft encounters with Saturn (November 13, 1980); these maneuvers involved rotating the spacecraft simultaneously about two or three axes while maintaining accurate pointing of the scan platform. Titan and Saturn earth occulation experiments and a ring scattering experiment are described. Target motion compensation and the effects of celestial sensor interference are also considered. Failure of the CST, which required an extensive reevaluation of the star reference and attitude control mode strategy, is discussed. Results analyzed thus far show that the system performed with high accuracy, gathering data deeper into Saturn's atmosphere than on any previous planetary encounter.