• CIVIL APPLICATION OF DIFFERENTIAL GPS

      Cardall, John D.; Cnossen, Richard S.; Magnavox Advanced Products and Systems Company (International Foundation for Telemetering, 1981-10)
      GPS has the potential of satisfying worldwide and local civil navigation requirements for Area Navigation (RNAV), Landings and Takeoffs under minimum ceilings and Advanced Air Traffic Control (ATC) Operations. Use of GPS in a differential mode in local areas is a key to achievement of this potential. This report describes the GPS system and its status; discusses GPS signal availability for the civil community; defines alternative differential GPS concepts; shows predicted performance enhancement achievable with differential GPS and the operational improvements which are expected.
    • BUILDUP AND REPLACEMENT OF NAVSTAR GLOBAL POSITIONING SYSTEM AND THE 18-SATELLITE CONSTELLATION

      Kruh, Pierre; THE AEROSPACE CORPORATION (International Foundation for Telemetering, 1981-10)
      Changes in the Navstar Global Positioning System (GPS) program in 1980 resulted in a reduction in the number of satellites from 24 to 18, and consequently led to a reconsideration of the original orbital configuration and a revision in the way buildup and replacement would be performed. During the buildup phase of the program, the Space Shuttle will be used, launching up to two Navstar satellites per mission on launches shared with other payloads and up to four satellites on dedicated launches. The original concept for the 24 satellite Navstar GPS constellation consisted of three orbit planes with eight uniformly distributed satellites per plane. In that configuration, all of the satellites on each Shuttle launch would be placed in only one plane. With the reduction from 24 to 18 satellites, it was determined that there were advantages to evaluating other constellations; rather than simply reducing the baseline three-plane, 24-satellite configuration to 18 satellites by removing two satellites from each plane, a larger number of planes may be preferred for the restructured program. Indeed, the current baseline configuration is a sixplane, 18-satellite constellation with three satellites per plane, and the three-plane constellation is an alternate configuration. Buildup and replacement studies for the new configurations that have been investigated have not only addressed the performance goals of Navstar GPS but have considered the economic use of the Shuttle as a launch platform. In addition, the launch constraints imposed by the Shuttle must be considered in the strategy used for buildup and replacement. This paper discusses the fundamentals of the buildup and replacement and the performance of the GPS with 18 satellites. The constellations discussed are the sixplane GPS baseline and the alternate three-plane configuration.
    • APPLICATIONS OF GPS PHASE COHERENCY

      Martin, Edward H.; Collins Government Avionics Division (International Foundation for Telemetering, 1981-10)
      The capabilities inherent in the Global Positioning System carrier and code modulation waveform provide an extremely precise and coherent spaceborne signal which may be applied to a variety of applications beyond the classical function of navigation. A description of rnechanizations based on both short and long baseline interferometry is given leading to applications for space vehicle attitude determination, azimuth bearing and level estimation, and precise survey location. Utilization of carrier phase changes is examined for dynamic vehicle surveys to provide a capability for accurately measuring gravity deflections in a real-time operational data processing system.
    • ULTRA LIGHTWEIGHT, LOW COST, TELEMETRY TRACKING SYSTEM

      Sullivan, Arthur; Electro Magnetic Processes, Inc. (International Foundation for Telemetering, 1981-10)
      Because of limited budgets, many telemetry applications can not be performed on a real time basis. Consequently, there is a need for a very low cost tracking system. In addition to being inexpensive, the system should be lightweight to minimize building or tower modifications required for installation, and to lower shipping and handling costs. In order to reduce weight and cost, EMP, Inc. has designed an aerodynamically smooth single axis tracking system with a multimode antenna, constructed almost entirely of graphite and utilizing a minimum number of parts. Using this material, the system will be as strong as a conventional system, but will be one half the weight and will be almost temperature insensitive. Furthermore, with the construction technique selected, considerable savings are realized in fabrication costs. Additionally, the step -track technique via a microprocessor controller was selected to eliminate the expensive autotracking feed and all the associated electronic circuitry required for high performance angle tracking. Since a single axis tracking system with a multi-mode antenna can cover a wide variety of missions, an elevation tracking axis is not required. In-close and near overhead passes are covered by the low gain antenna. Switching between the antennas is accomplished automatically based on received signal strength.
    • MULTIBAND OMNIDIRECTIONAL TELEMETRY ANTENNA

      Johnson, Russ; Metzler, Tom; Ball Aerospace Systems Division (International Foundation for Telemetering, 1981-10)
      With the increasing sophistication of telemetry and tracking systems, an omnidirectional antenna plays a major role in assuring adequate telemetry system signal to noise ratio regardless of test platform orientation. However, the increased demands on antenna performance often impact the antenna complexity, size and weight. This paper describes a simple yet extremely rugged antenna designed to conformally mount to a large diameter missile and provide omnidirectional coverage at four discrete frequency bands while minimizing structural impact on the missile.
    • “MINITRACKER” A PORTABLE S-BAND AUTOTRACK ANTENNA

      Skach, Leonard J.; Oklahoma State University Electronics Laboratory (International Foundation for Telemetering, 1981-10)
      The “Minitracker” is designed to be a highly portable S-Band autotrack antenna system, to be used to receive telemetry data from sounding rockets launched at remote launch sites when fixed-site telemetry trackers are not available. “Minitracker” uses a single channel mono-pulse RF feed with a four foot parabolic reflector. The tracker can be disassembled into small subassemblies and packed in plastic cases which can easily be shipped by commercial air freight. The controls consist of a 8 3/4" (22.225 cm) X 19" (48.26 cm) X 19" (48.26 cm) control console and an S-Band Receiver. The OSU “TRADAT” system can be used with “Minitracker” to provide trajectory data. This report will consist of a “Minitracker” system description which will include systems performance specifications and simplified circuit descriptions. The operation procedures section will consist primarily of aids in autotrack acquisition of target vehicles, and a brief description of information available when used in conjunction with the OSU built “TRADAT” system.
    • A MICROSTRIP ANTENNA FEED FOR A PARABOLOIDAL REFLECTOR WITH SIMULTANEOUS RCP AND LCP POLARIZATION

      Post, Cecil C.; New Mexico State University (International Foundation for Telemetering, 1981-10)
      A microstrip antenna feed developed for a paraboloidal reflector provides simultaneous right circular and left circular polarization and 20 dB sidelobe levels over moderate bandwidths. This arrangement obviates the need for the usual waveguide orthomode transducer and thus decreases the weight of the feed. The basic feed element is a square patch .558λ x .558λ fed at two orthogonally situated coaxial feed points inset 0.118λ from the edge, as described by Millar and Carver (Proceedings 1980 University of Illinois Allerton Antennas Application Symposium).
    • END COUPLED PARASITIC MICROSTRIP ANTENNA

      KALOI, MO; PACIFIC MISSILE TEST CENTER (International Foundation for Telemetering, 1981-10)
      A Parasitic Microstrip Antenna Array is discussed. The array consists of different lengths of microstrip radiating elements spaced apart in an end-to-end arrangement. Only one element is actively fed at its feedpoint, and energy emanating from the fed element is primarily coupled to parasitic elements by the electric field generated in the fed element. The radiating pattern is determined by the phase relationship and amplitude distribution between the excited fed element and the parasitic elements. This antenna configuration exhibits higher end fire gain along the missile axis than obtained with arrays having elements individually fed.
    • THE USE OF THE CONICAL SCAN EARTH SENSOR IN COMMUNICATION SATELLITE APPLICATIONS

      Fowler, Robert Z.; ITHACO, Inc. (International Foundation for Telemetering, 1981-10)
      Infra-red horizon sensors are almost universally used as the primary attitude sensor for pitch and roll on present day three-axis stabilized communication satellites. When used with a momentum wheel, yaw is also controlled without direct sensing. The application flexibility of the mechanically scanned Conical Earth Sensor, and it’s recent availability as a component designed for precision, long life performance have resulted in renewed interest in its use on communication satellites. The Conical Earth Sensor will provide accurate on-orbit attitude sensing in pitch and roll. It can provide attitude sensing all the way from the shuttle orbit to synchronous for booster control, and is particularly attractive for multiple burn, multiple orbit transfer. It can provide accurate nadir sensing 100% of the time in the highly elliptical Molniya twelvehour orbit. It can facilitate wide angle attitude sensing for antennae calibration maneuvers. It can be used in a static mode as a horizon crossing indicator for spacecraft that go up as spinners, and then for normal on-orbit sensing as a scanner. It can be readily hardened to both nuclear and lazer threats, unlike static sensors that are highly susceptible to thermal transients. It has a simple, rugged, and stable construction that is not sensitive to resonance effects from other mechanical devices on the spacecraft such as momentum or reaction wheels.
    • TWO AXIS CLOSED-LOOP ANTENNA POINTING FOR A DUAL-SPIN SPACECRAFT

      Smay, John W.; Hughes Aircraft Company (International Foundation for Telemetering, 1981-10)
      This paper describes the control and sensing techniques and practical implementation used to obtain precision antanna pointing on a class of commercial communication satellites. The basic spacecraft bus is a dual-spin gyrostat with momentum of order 1500 ft-lb-sec. Spin is about a minimum axis of inertia and active damping using the despin motor and platform product of inertia is employed for nutation stabilization. Using two axis RF beacon tracking, steady state pointing accuracy exceeding 0.025° (3σ) in roll and pitch and 0.1° (3σ) in yaw is achieved. This accuracy is approached during orbit and attitude trim thrusting maneuvers as well.
    • TDRS ANTENNA AUTOTRACK LOOP

      Schmeichel, Harry; TRW Defense and Space Systems Group (International Foundation for Telemetering, 1981-10)
      The Tracking Data and Relay Satellite (TDRS) has two large, gimballed antennas which will relay information between earth-orbiting satellites and a ground terminal in New Mexico at data rates up to 300 million bits per second. This relay service requires closedloop tracking of user satellites at K-band frequencies with a pointing accuracy of 0.06°. An autotrack loop, closed through a ground-based computer, performs this RF beam pointing function for each single-access (SA) antenna. The autotrack system basically consists of two stepper motors to move the antenna, an onboard RF monopulse system to sense the pointing error and command generation equipment on the ground to close the loop. It is shown how system models and observations are combined to stabilize and improve the pointing performance of this lowbandwidth, closed-loop tracking system. Antenna pointing performance is demonstrated by simulation.
    • DSCS-III ATTITUDE CONTROL SYSTEM

      Bonello, D.; Basuthakur, S.; Valley Forge Space Center (International Foundation for Telemetering, 1981-10)
      The DSCS-III (Defense Satellite Communications System) Spacecraft was designed and built for the Air Force Space Division by General Electric Space Division in Valley Forge, PA. Development of this satellite started in 1978 and was culminated in the recent (May 81) completion of testing of the first flight unit. The attitude control system for this synchronous orbit spacecraft is a three-axis zero momentum, general microprocessor controlled concept that not only provides attitude and velocity control during the normal seven year orbital life, but also provides provisions to operate and maintain control during special circumstances such as failed battery eclipses, lunar eclipses, and nuclear events. In addition, the attitude control system electronics and embedded software system provides the capability to drive the single axis solar array, two axis gimbal dish antenna, and translates ground commands into beam pattern reconfiguration driver signals for the phase shifters and variable power dividers of the payload multiple beam antennas. The control system equipment compliment consists of a redundant passive radiation balance earth sensor, solar array yoke mounted analog sun sensors, and a yaw rate gyro as the sensing elements, the aforementioned general purpose microprocessor (Attitude Control Electronics containing 8K of PROM memory and 1K of RAM in which is implemented the control logic and algorithms, four skewed reaction wheels for normal orbital control torquing/momentum storage and 16 one-pound hydrazine thrusters for initial acquisition and orbit adjust maneuvers and wheel unloading. The basic requirements to which this system was designed are to (i) acquire an earth pointing reference from arbitrary initial attitudes and rates of 1.1°/sec per axis, maintain control during initial inclination error removal (maximum of 2. 5 degrees) to within 1, 1 and 2 degrees for the roll, pitch, and yaw axes respectively for all times of year and orbit positions, (ii) maintain pitch, roll, and yaw errors to less than 0.08, 0.08, and 0.8 degrees during normal orbital operations (iii) maintain orientation of the solar array to within 1° of the sunline, (iv) establish and control station latitude and longitude to within + 0.1°, (v) provide the capability to recover from the effects of a nuclear event via autonomous detection and corrective action, (vi) provide the capability to reorient/reconfigure the payloac gimballed dish antenna and multiple beam antennas and (vii) provide the capability to modify up to 1K of the PROM stored software using a ground commanded mode. This DSCS-III ACS as designed and tested meets all of its requirements with a system weighting only 83. 3 pounds and using approximately 64 watts of power.
    • A COMPUTER-AIDED SIMULATION PROGRAM DESIGN FOR TESTING THE FLIGHT CONTROLLER OF A THREE AXIS STABILIZED SPACECRAFT

      Basuthakur, Sibnath; General Electric Company (International Foundation for Telemetering, 1981-10)
      To insure a reliable performance of any spacecraft over its long mission life, a thorough and coordinated attitude control subsystem testing must be conducted. The three axis motion Simulator-Hybrid computer facility at General Electric has provided the capability of testing the Attitude Control Electronics (ACE) for various satellite programs including Japanese satellite program BSE and Defense Communication Satellite DSCS-III. Although the facility has provided complete verification of analysis and simulation of all operating modes in a closed-loop fashion, the checkout procedure has proven to be extremely timeconsuming. It requires real time dedicated computer support. In addition, limited sensor field of view may, in some instances, limit the scope of the test. The objective of this paper is to underline an alternate philosophy of the subsystem testing that has been extensively used to qualify the DSCS-III flight control system under various environments. It is designed to compare, on a bit by bit basis, all critical controller internal and output parameters between the flight control algorithms embedded in the ACE and a validated simulator controller. The simulated controller (truth model) is validated after careful analyses and simulation of all operating modes under all possible initial conditions. All controller parameters to be compared are assigned to CPU test port and the telemetry port. This computer-aided testing program is used to process CPU output data in an off-line autonomous basis to validate the control algorithms embedded in the ACE.
    • ACCURATE ANTENNA REPOINTING FOR PATTERN MAPPING

      HOLTZ, L. VAN; European Space Research & Technology Center (International Foundation for Telemetering, 1981-10)
      The attitude control system of the European Space Agency’s Orbital Test Satellite (OTS) was originally designed for nominal earth pointing with only limited bias capability (of up to 2.5° in pitch and roll) which is more than adequate to remove earth sensor/wheel misalignments that could be incurred during or following launch. Subsequently, a need was expressed to support off-nominal coverage missions and antenna mapping tests. A method was thus defined that would provide greater repointing capability (up to 4.6°) while retaining the accuracy available with the precision infrared sensor. The method is outlined, and the repointing limitations indicated that are inherent to the OTS sensor design (13° in roll and 7° in pitch). Operations and conditions are stated that enable these extremes to actually be reached. The error budget is presented for the case of antenna mapping, demonstrating that attitude restitution can be made so that beam centre position can be determined to an accuracy of 0.1° half cone angle. The significant advantages of the described method are that only one ground station is required, and that results can be available within 24 hours following completion of the test. Results obtained with OTS are referred to, that support the claims. Finally desirable design modifications are discussed that could allow further increases in repointing capability of future satellites.
    • SCIENCE AND APPLICATIONS SPACE PLATFORM COMMUNICATIONS AND DATA MANAGEMENT SYSTEM

      Kasulka, L. H.; McDonnell Douglas Astronautics Company (International Foundation for Telemetering, 1981-10)
      The development of space platforms represents the next logical step in the exploration and utilization of space. Such platforms promise cost-effective means for performing both scientific and applications missions, such as surveys of Earth resources, for example, in low Earth orbit. Payloads mounted on these platforms can perform missions for longer periods of time than are currently available to payloads mounted in the Shuttle’s payload bay. In addition, these platforms can provide a variety of services, including a centralized power source, command and data acquisition, communications, pointing and environmental control, as well as periodic Shuttle visits for performing maintenance tasks, replenishing consumables, and replacing payloads. These platforms must be able to provide data and communications services to groups of payloads consisting of individual payloads that may or may not have common objectives and operating characteristics, and where the payload mix on a platform changes periodically during the orbital life of the platform. Appropriate data systems can be provided to support a platform development program and modest extensions of existing technology will allow these platforms to accommodate the evolution of payloads foreseen through the 1980’s.
    • INTELSAT V SPACECRAFT TELEMETRY COMMAND AND RANGING

      Johnson, Charles E.; Communications Satellite Corporation Palo Alto (International Foundation for Telemetering, 1981-10)
      The INTELSAT V communications satellite was designed and assembled by the Ford Aerospace and Communications Corporation of Palo Alto, California, for the International Telecommunications Satellite Consortium (INTELSAT). The Communications Satellite Corporation (COMSAT) is the designated United States representative to INTELSAT and also performs technical service for INTELSAT in monitoring the design, fabrication, and test of communications satellites. The TT&C subsystem consists of two functionally redundant and independent command and telemetry channels, the major elements of which are shown in Figure 1. The telemetry subsystem provides two data channels for formatting and transmitting data received from sensors, transducers, and status indicators in the various subsystems of the spacecraft. In addition, the output of a command receiver can be connected to a telemetry transmitter to form a ranging transponder. The telemetry unit can provide normal or dwell PCM data in NRZ-M format modulated on a 32 KHz subcarrier. The telemetry transmitter phase modulates one of the data subcarriers or ranging tones on a 4 GHz band downlink carrier. The transmitter output is routed directly to an earth coverage conical horn for transmission at a level of approximately 0.0dbW. The transmitter output can also be routed through a zone communications channel TWTA to a telemetry omni-directional bicone antenna for extended coverage. Commands and ranging tones are received on a 6 GHz band uplink carrier through dual Omni pattern antennas. The received signal is routed through a passive filter to the two command receivers where the frequency modulated command or ranging tones are detected. The command tones are routed to the command units for processing. The command transmission is either a command message, consisting of 58 serial bit, or a command execute. The command message includes the address of the command unit which is to be used and what specific command function is to be executed. The command units provide the capability for pulse, discrete relay, and proportional relay command functions required by the various subsystems of the spacecraft.
    • GEOSTATIONARY OPERATIONAL ENVIRONMENTAL SATELLITE (GOES)

      Nakamura, Apryll M.; Mallette, Leo A.; Hughes Aircraft Company (International Foundation for Telemetering, 1981-10)
      The GOES satellites are multifunctional satellites whose primary function is to provide continuous measurements of the earth’s surface and atmosphere from two geostationary orbit locations: 75°W and 135°W. This objective is accomplished with the Visible infrared spin scan radiometer Atmospheric Sounder (VAS). The atmospheric sounder is a new feature which will add a third dimension to the photographs of the earth seen nightly by TV newscast viewers. The satellite also contains a Space Environment Monitor (SEM) which includes three instruments: a magnetometer, a solar X-ray sensor and an energetic particle sensor (EPS), which monitor solar flares and near earth space environment. The satellite contains a communications system which, in addition to transmitting VAS, SEM and housekeeping data to earth, provides relay capabilities for the stretched VAS and weather facsimile (WEFAX) data, as well as for the Data Collection Platform (DCP). A sketch of the satellite is given in Figure 1. The telemetry system encompasses two subsystems: RF communications and baseband assemblies. A general diagram of the telemetry system is shown in Figure 2. The telemetry system consists of two functional operations; real time telemetry data which is frequency modulated into IRIG 12 and IRIG B, and PCM data which is phase modulated. The baseband assemblies collect, format and modulate the data. The RF system sums the PCM and IRIG signals, phase modulates it onto both the CDA and STDN carriers, and transmits it to the ground stations. The following sections will describe in further detail the operation of the PCM and real time telemetry functions, and a description of the satellite RF communications system.
    • EVOLUTION OF TELEMETRY AND COMMAND SYSTEMS FROM EARLY BIRD TO INTELSAT V

      Magnusson, S. Erland; International Telecommunications Satellite Organization (International Foundation for Telemetering, 1981-10)
      The telemetry and command system on INTELSAT satellites has gone through an evolution from the early series of satellites where simplicity and satellite reliability was emphasized to the latest series of satellites where communications systems reliability is emphasized. The early telemetry and command systems were integrated with the communications subsystem on the satellite and had some redundancy. Later systems are autonomous with complete redundancy and cross-strapping between the systems to a large extent.
    • COMMUNICATIONS FOR THE ESA GIOTTO (COMET HALLEY ENCOUNTER) MISSION

      Wasse, Michael P.; European space Technology Centre (International Foundation for Telemetering, 1981-10)
      This space mission will investigate the structure of the comet by passing close to the nucleus. The spacecraft will first be injected into a geostationary transfer orbit where a perigee boost motor will deliver the kick necessary to encounter the comet post perihelion, as it passes through the ecliptic plane at a distance of 0.98 AU from the earth. The spinning spacecraft and the use of a shield for protection from the dust present in the comet atmosphere dictate the use of a despun high gain antenna with inclined beam. The telemetry downlink at 40 KBps is in X Band and will be received on the Parkes (Australia) radio telescope which will be specially adapted for the task. The trade offs leading to the selected communications configuration are described along with the various spacecraft hardware items such as antennas, transponder, decoder and twta.
    • COLOCATED COMSTARS

      Lee, David J.; Guthrie, W. Coleman; McKee, Walter S., Jr.; COMSAT General Corporation (International Foundation for Telemetering, 1981-10)
      Both AT&T and GSAT are presently using the COMSTAR to expand and diversify their domestic public dial telephone networks. COMSAT General’s TTC&M earth station facilities at Southbury, Connecticut, and Santa Paula California, continuously monitor the status of the COMSTARs. For the purpose of increasing the likelihood of availability and maximizing the number of communications transponders to the users, at the end of the spacecraft design lifetime, COMSAT General was able to reach a business agreement with AT&T and obtain FCC authorization to launch COMSTAR D-4 and colocate D-1 and D-2. Therefore, as far as the communications earth stations are concerned, colocated D-1 and D-2 with a coordinated 24 TWT configuration are technically identical to a single COMSTAR satellite. There is, however, a significant increase in satellite lifetime due to reduced solar array and battery loading (about half) since each satellite now operates with twelve TWTs on instead of the usual twenty-four for a COMSTAR. This paper will describe the TTC&M earth stations’ modification to accommodate the colocated satellites. Operational considerations and some actual operational experience will also be discussed.