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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Inflight redesign of the IUE attitude control system

The one- and two-gyro system designs of the International Ultraviolet Explorer (IUE) attitude control system (ACS) are examined. The inertial reference assembly that provides the primary attitude reference for IUE consists of six rate sensors which are single-axis rate integrating gyros. The gyros operate in a pulse rebalanced mode that produces an output pulse for 0.01 arcsec of motion about the input axis. The functions of the fine error sensor, fine sun sensor (FSS), the IUE reaction wheels, the onboard computer, and the hold/slew algorithm are described. The use of the hold/slew algorithm to compute the control voltage for the ACS based on the Kalman filter is studied. A two-gyro system was incorporated into IUE following gyro failure. The procedures for establishing attitude control with the two-gyro design based on the FSS is analyzed. The performance of the two-gyro system is evaluated; it is observed that the pitch and yaw gyro control is 0.24 arcsec and the control is sufficient to permit extended periods of observation.

Femiano, M. D.↗
NASA NTRS2007

Fault Detection and Correction for the Solar Dynamics Observatory Attitude Control System

The Solar Dynamics Observatory is an Explorer-class mission that will launch in early 2009. The spacecraft will operate in a geosynchronous orbit, sending data 24 hours a day to a devoted ground station in White Sands, New Mexico. It will carry a suite of instruments designed to observe the Sun in multiple wavelengths at unprecedented resolution. The Atmospheric Imaging Assembly includes four telescopes with focal plane CCDs that can image the full solar disk in four different visible wavelengths. The Extreme-ultraviolet Variability Experiment will collect time-correlated data on the activity of the Sun's corona. The Helioseismic and Magnetic Imager will enable study of pressure waves moving through the body of the Sun. The attitude control system on Solar Dynamics Observatory is responsible for four main phases of activity. The physical safety of the spacecraft after separation must be guaranteed. Fine attitude determination and control must be sufficient for instrument calibration maneuvers. The mission science mode requires 2-arcsecond control according to error signals provided by guide telescopes on the Atmospheric Imaging Assembly, one of the three instruments to be carried. Lastly, accurate execution of linear and angular momentum changes to the spacecraft must be provided for momentum management and orbit maintenance. In th~sp aper, single-fault tolerant fault detection and correction of the Solar Dynamics Observatory attitude control system is described. The attitude control hardware suite for the mission is catalogued, with special attention to redundancy at the hardware level. Four reaction wheels are used where any three are satisfactory. Four pairs of redundant thrusters are employed for orbit change maneuvers and momentum management. Three two-axis gyroscopes provide full redundancy for rate sensing. A digital Sun sensor and two autonomous star trackers provide two-out-of-three redundancy for fine attitude determination. The use of software to maximize chances of recovery from any hardware or software fault is detailed. A generic fault detection and correction software structure is used, allowing additions, deletions, and adjustments to fault detection and correction rules. This software structure is fed by in-line fault tests that are also able to take appropriate actions to avoid corruption of the data stream.

Starin, Scott R.↗
NASA NTRS1986

Cosmic Background Explorer (COBE) transfer orbit attitude control system

The Cosmic Background Explorer (COBE) spacecraft will be launched by the Shuttle from Vandenberg AFB into a 300 km altitude, 99 deg inclination, 6 a.m. or 6 p.m. ascending node orbit. After release from the Remote Manipulator System (RMS) arm, an on-board monopropellant hydrazine propulsion system will raise the orbit altitude to 900 km. The spacecraft continuously spins during transfer orbit operations with the spin axis nominally horizontal and in or near the orbit plane. The blowdown propulsion system consists of twelve 5 lb thrusters (3 'spin', 3 'despin', and 6 'axial') with the latter providing initially 30 lb of force parallel to the spin axis for orbit raising. The spin/despin jets provide a constant roll rate during the transfer orbit phase of the mission and the axials control pitch and yaw. The axial thrusters are pulsed on for attitude control during coast periods and are normally on- and off-modulated for control during orbit raising. Attitude sensors employed in the control loops include an array of two-axis digital sun sensors and three planar earth scanners for position measurements, as well as six gyroscopes for rate information. System redundancy is achieved by means of unique three-axes-in-a-plane geometry. This triaxial concept results in a fail-safe operational system with no performance degradation for many different component failure modes.

Placanica, Samuel J.↗
NASA NTRS1987

Using a horizon sensor for Mars observer spacecraft attitude control

A modified earth horizon sensor design is proposed for the Mars Observer mission. The horizon sensor modification are based on Mars horizon radiance profiles. Attention is given to the Mars Observer's attitude control system, whose primary mode is horizon sensor-based and therefore obviates the use of an onboard spacecraft ephemeris; control accuracy of the order of 10 mrad 3sigma/axis is achieved. The attitude control system's second, celestial-backup mode, wich requires greater operational effort, employs an onboard celestial sensor as the spacecraft ephemeris; its control accuracy is of the order of 8 mrad 3sigma/axis.

Langmaier, J. K.↗
NASA NTRS1973

Model-reference attitude control and reaction control jet engine placement for space shuttle

Analytical studies on the theoretical aspects of thrust vector control of large space vehicles were conducted. A system for attitude control of the space shuttle vehicle was developed. Major accomplishments of the project are: (1) investigation of a model-reference adaptive control scheme for controlling the space shuttle attitude and (2) determination of optimum placement of reaction control jet engines on space shuttles.

Boland, J. S., III↗
NASA NTRS1979

Attitude control requirements for future space systems

Landsat-D and the Large Space Telescope represent current state-of-the-art systems with precise requirements placed on attitude control. Future systems for planetary stations, high precision earth monitoring and large precision deployable and erectable platforms project still more severe constraints and requirements on attitude control, including the requirement for many enabling and highly enhanced technologies beyond current state of the art. Two trend projections are identified for the areas of (1) precision pointing systems for earth orbiters and planetary spacecraft and (2) onboard high-capacity fast controllers for distributed control systems.

Dahlgren, J. B.↗
NASA NTRS1973

Precise attitude control of the Stanford relativity satellite.

A satellite being designed by the Stanford University to measure (with extremely high precision) the effect of General Relativity is described. Specifically, the satellite will measure two relativistic precessions predicted by the theory: the geodetic effect (6.9 arcsec/yr), due solely to motion about the earth, and the motional effect (0.05 arcsec/yr), due to rotation of the earth. The gyro design requirements, including the requirement for precise attitude control and a dynamic model for attitude control synthesis, are discussed. Closed loop simulation of the satellite's natural dynamics on an analog computer is described.

Bull, J. S.↗
NASA NTRS1975

Launch vehicle attitude control

The minor loop and minor loop support modules which maintain correct vehicle attitude control by generating vehicle attitude error signals were verified. The logic is checked during boost and orbit (unusual timing situations necessitate the additional orbit checks) and the constants used by these modules checked at every point during the flight. Since the same logic instructions are used throughout the flight with only the constants changing, verification of the logic at two points and the constants at all points verify the logic throughout the flight.

Source record↗
NASA NTRS1974

An elementary magnetic attitude control system

A three-axis closed-loop attitude-control system for earth-observatory momentum-bias spacecraft is proposed. A horizon scanner and magnetometer are employed to measure attitude errors; no yaw sensor is required. Appropriate control signals are generated and used to command variable-strength electromagnets, which interact with the geomagnetic field to reduce observed errors. This system provides for initial acquisition, precession control, nutation damping, and pitch-axis momentum control. All functions are performed autonomously - no ground-station interaction is required. Analytical expressions predicting system response are compared with numerical solutions of the governing equations, and with the results of the application of Floquet Theory.-

Stickler, A. C.↗
NASA NTRS2018

Distributed Attitude Control and Maneuvering for Deep Space SmallSats

The aim of the Distributed Attitude Control and Maneuvering for Deep Space SmallSats project is to advance a multi-purpose, deep space mission-enabling technology for low-power attitude and thermal control of small satellites to a flight demonstration technology readiness level (TRL). The film-evaporation microelectromechanical systems tunable array (FEMTA) small satellite technology combines innovative microelectromechanical systems (MEMS) microfabrication and microscale effects in fluid surface tension to produce a thermally actuated capillary valve. Using water as the propellant, the FEMTA thruster can generate finely controllable thrust at a thrust to power ratio of about 200 microNewton per Watt (W).

thermal control↗
NASA NTRS1998

An approach to the design and implementation of spacecraft attitude control systems

Over 39 years and a long list of missions, the guidance, navigation, and control (GN&C) groups at the Goddard Space Flight Center have gradually developed approaches to the design and implementation of successful spacecraft attitude control systems. With the recent creation of the Guidance, Navigation, and Control Center at Goddard, there is a desire to document some of these design practices to help to ensure their consistent application in the future. In this paper, we will discuss the beginnings of this effort, drawing primarily on the experience of one of the past attitude control system (ACS) groups at Goddard (what was formerly known as Code 712, the Guidance, Navigation, and Control Branch). We will discuss the analysis and design methods and criteria used, including guidelines for linear and nonlinear analysis, as well as the use of low- and high-fidelity simulation for system design and verification of performance. Descriptions of typical ACS sensor and actuator hardware will be shown, and typical sensor/actuator suites for a variety of mission types detailed. A description of the software and hardware test effort will be given, along with an attempt to make some qualitative estimates on how much effort is involved. The spacecraft and GN&C subsystem review cycles will be discussed, giving an outline of what design reviews are typically held and what information should be presented at each stage. Finally, we will point out some of the lessons learned at Goddard.

ODonnell, James R., Jr.↗