SEARCH · Search NASA

Results for “INERTIAL GUIDANCE”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 73 records · Page 4

All sources NASA NTRS DOE OSTI

NASA NTRS1972

Ultra-precise attitude test facility for flight spacecraft

A test facility was designed and built for conducting complete dynamic systems tests of the OAO spacecraft. The operating principle of the facility is illustrated for a single control axis. The test facility is described which was built and used to test the stabilization and control system of the OAO B spacecraft. Fine pointing of the spacecraft was demonstrated to better than 0.5 seconds of arc. Other spacecraft functions tested were gimbaled and boresight star tracker operations, coarse and fine slew maneuvers, inertial guidance operation, settling time of the spacecraft control system, and transfers between all modes of control.

Devine, E. J.↗

NASA NTRS1971

The Delta and Thor/Agena launch vehicles for scientific and applications satellites.

Description of the Delta Model 904 and the Thor/Agena Model 9A4 scientific and applications satellite launch vehicles, with projections of future growth and launch costs. These launch vehicles are shown to offer scientific and applications satellite mission planners a broad spectrum in performance capabilities together with unprecedented mission flexibility. Depending on the mission, these two medium class launch vehicles can be configured on the new universal boattail (UBT) Thor booster in either two or three stages with thrust augmentation of the UBT ranging from three to nine strap-on solid propellant motors. Both vehicles incorporate strapdown inertial guidance systems that allow flexible mission programming by computer so ftware changes rather than by adjustments.

Gunn, C. R.↗

NASA NTRS1973

The Delta launch vehicle Model 2914 series

Description of a new, medium-class Delta launch-vehicle configuration, the three-stage Model 2914. The first stage of this vehicle is composed of a liquid-propellant core which is thrust-augmented with up to nine strap-on solid-propellant motors. The second stage, recently uprated with a strap-down inertial guidance system, is now being modified to adapt the liquid-propellant descent engine from the Apollo Lunar Excursion Module. The third stage is a spin-stabilized solid-propellant motor. The Model 2914 is capable of injecting 2040 kg into low earth orbit, 705 kg into geosynchronous transfer orbit, or 455 kg into an escape trajectory.

Gunn, C. R.↗

NASA NTRS1975

Space Shuttle flight control system

The Space Shuttle is a control stabilized vehicle with control provided by an all digital, fly-by-wire flight control system. This paper gives a description of the several modes of flight control which correspond to the Shuttle mission phases. These modes are ascent flight control (including open loop first stage steering, the use of four computers operating in parallel and inertial guidance sensors), on-orbit flight control (with a discussion of reaction control, phase plane switching logic, jet selection logic, state estimator logic and OMS thrust vector control), entry flight control and TAEM (terminal area energy management to landing). Also discussed are redundancy management and backup flight control.

Klinar, W. J.↗

NASA NTRS1976

Study of Arctic sea ice drift from L-band synthetic aperture radar

As part of the Arctic Ice Dynamics Joint Experiment (AIDJEX) several repetitive coverages of L-band (25 cm wavelength) side-looking airborne radar images have been flown over coastal areas of Alaska and a test area in the Arctic. These images allow the analysis of sea ice and its drift. Radar is particularly suited for the mapping and interpretation of Arctic sea ice due to independence from sunlight and the capability to penetrate clouds. Ice floes and leads can be readily identified on the radar images. Measurement of ice floe drift is based on the transformation of radar image coordinates into a geocentric coordinate system using inertial guidance data from the survey aircraft. The paper will demonstrate an example of Arctic ice drift measurements from L-band synthetic aperture radar imagery with an absolute accuracy of about 5%. The conclusions are of particular value in view of planned spaceborne side-looking radar missions in polar orbits.

Leberl, F.↗

NASA NTRS1978

Optical gyroscope

Instrument uses phase difference between two beams of light to measure rotation. It is considerably simpler and more reliable than conventional spinning-mass gyroscopes used for inertial guidance and is more compact, lighter, and potentially less expensive. Moreover, optical gyroscope requires no warmup period. Although conceived for spacecraft and satellite stabilization, gyroscope should also find applications in flight instruments for private, commercial, and military aircraft.

Goldstein, R. M.↗

NASA NTRS1980

Space applications of superconductivity - Instrumentation for gravitational and related studies

Superconducting instruments for gravitational studies and their state of development are discussed. Consideration is given to accelerometers, gravimeters, superconducting gyroscopes, and inertial guidance and navigation systems. This is followed by discussion of applications to tests of gravitational theories, determination of elastic properties of planets, and navigation.

Peterson, R. L.↗

NASA NTRS1981

NASA to launch second business communications satellite

The two stage Delta 3910 launch vehicle was chosen to place the second small business satellite (SBS-B) into a transfer orbit with an apogee of 36,619 kilometers and a perigee of 167 km, at an inclination of 27.7 degrees to Earth's equator. The firing and separation sequence and the inertial guidance system are described as well as the payload assist module. Facilities and services for tracking and control by NASA, COMSAT, Intelsat, and SBS are outlined and prelaunch operations are summarized.

Source record↗

NASA NTRS1981

OAO-3 end of mission tests report

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.

Kalil, F.↗

NASA NTRS1984

Doppler radiation sensing of Shuttle angle of attack and TAS during entry

Space Shuttle true airspeed, angle of attack, and sideslip angle are currently derived from inertial guidance information. A new method is proposed which offers a potential improvement in Shuttle safety during entry. Angle of attack, sideslip angle and true airspeed could be measured directly at heights from 120 km down to 20 km by Doppler sensing three independent true airspeeds along the Shuttle body axes. Two types of Doppler measurement sensors, employing either passive detection of atmospheric radiation or coherent detection of scattered laser light are discussed. The proposed technique is essentially solid-state and robust, and is well suited for use in future small hypersonic vehicles that require flight control in the Upper Atmosphere of the earth or in probes destined for the other planets.

Foale, C. M.↗

NASA NTRS1984

Coupled orbiting inertial reference systems and Geopotential Research Mission (GRM) geodesy

Two spacecraft, one following the other with a separation of a few hundred kilometers and orbiting at an altitude of 160 km in a 90-degree polar orbit, are used to detect the minute variations in the earth's gravitational field. Housed in a capacitive cavity located at the centroid of the spacecraft is a free-floating metallic spherical proof-mass, which is the sensor controlling the thrusters that nullify all nongravitational forces on the spacecraft. The proof-mass in each spacecraft is the inertial-guidance reference for its host spacecraft, and the proof-masses are coupled together by a millimeter-wave link. This coupled system detects minute changes in the spacecraft's orbital positions and thus measures the anomalies present in the gravitational field of the earth. To derive a global geodetic model of the gravitational field with an amplitude resolution of 1 milligal and a spatial resolution of 100 km requires that the relative velocity between the proof-mass references be determined to within 1 micron/s.

Keating, T.↗

NASA NTRS1986

Development of a sensitive superconducting gravity gradiometer for geological and navigational applications

A sensitive and stable gravity gradiometer would provide high resolution gravity measurements from space. The instrument could also provide precision tests of fundamental laws of physics and be applied to inertial guidance systems of the future. This report describes research on the superconducting gravity gradiometer program at the University of Maryland from July 1980 to July 1985. The report describes the theoretical and experimental work on a prototype superconducting gravity gradiometer. The design of an advanced three-axis superconducting gravity gradiometer is also discussed.

Paik, H. J.↗

NASA NTRS1990

Characterizing performance of ultra-sensitive accelerometers

An overview is given of methodology and test results pertaining to the characterization of ultra sensitive accelerometers. Two issues are of primary concern. The terminology ultra sensitive accelerometer is used to imply instruments whose noise floors and resolution are at the state of the art. Hence, the typical approach of verifying an instrument's performance by measuring it with a yet higher quality instrument (or standard) is not practical. Secondly, it is difficult to find or create an environment with sufficiently low background acceleration. The typical laboratory acceleration levels will be at several orders of magnitude above the noise floor of the most sensitive accelerometers. Furthermore, this background must be treated as unknown since the best instrument available is the one to be tested. A test methodology was developed in which two or more like instruments are subjected to the same but unknown background acceleration. Appropriately selected spectral analysis techniques were used to separate the sensors' output spectra into coherent components and incoherent components. The coherent part corresponds to the background acceleration being measured by the sensors being tested. The incoherent part is attributed to sensor noise and data acquisition and processing noise. The method works well for estimating noise floors that are 40 to 50 dB below the motion applied to the test accelerometers. The accelerometers being tested are intended for use as feedback sensors in a system to actively stabilize an inertial guidance component test platform.

Sebesta, Henry↗

NASA NTRS1991

An autonomous rendezvous and docking system using cruise missile technologies

In November 1990 the Autonomous Rendezvous & Docking (AR&D) system was first demonstrated for members of NASA's Strategic Avionics Technology Working Group. This simulation utilized prototype hardware from the Cruise Missile and Advanced Centaur Avionics systems. The object was to show that all the accuracy, reliability and operational requirements established for a space craft to dock with Space Station Freedom could be met by the proposed system. The rapid prototyping capabilities of the Advanced Avionics Systems Development Laboratory were used to evaluate the proposed system in a real time, hardware in the loop simulation of the rendezvous and docking reference mission. The simulation permits manual, supervised automatic and fully autonomous operations to be evaluated. It is also being upgraded to be able to test an Autonomous Approach and Landing (AA&L) system. The AA&L and AR&D systems are very similar. Both use inertial guidance and control systems supplemented by GPS. Both use an Image Processing System (IPS), for target recognition and tracking. The IPS includes a general purpose multiprocessor computer and a selected suite of sensors that will provide the required relative position and orientation data. Graphic displays can also be generated by the computer, providing the astronaut / operator with real-time guidance and navigation data with enhanced video or sensor imagery.

Jones, Ruel Edwin↗

NASA NTRS1993

Miniature Optical Wide-Angle-Lens Startracker (Mini-OWLS)

This paper provides a brief overview of the design considerations and the current status of the Miniature Optical Wide-Angle Lens Startracker Program. Mini-OWLS offers a revolutionary alternative to the conventional startracker. It is a small, lightweight, low cost, high performance startracker that can be used in a variety of applications including calibration and alignment of Inertial Measurement Units (IMU's) Mini-OWLS makes use of a strap down design incorporating Holographic Optical Elements (HOES) in place of conventional optics. HOES can be multiplexed so that the same aperture can be used for multiple separate optical paths looking in several directions simultaneously without startracker rotation. Additionally, separate Schmidt corrector plates are not required to compensate for spherical aberration. The optical assembly, or what would normally be considered as the telescope, is less than 20 cc in volume, weighs less than 55 grams, and contains the equivalent of three individual telescopes. Each one has a 4 deg Field of View (FOV) with a field of regard of 48 square degrees. Mini-OWLS has a bandwidth of approximately 300 nm in or near the visible wavelength. The projected resolution of the startracker is 5 to 10 arcseconds, depending on the centroiding algorithm used. The Mini-OWLS program was initiated last year and represents a miniaturized version of a similar design for aeronautical applications. The contract is managed by Wright Laboratory, Air Force Systems Command, Wright-Patterson AFB, Ohio, with funding from the Strategic Defense Initiative Organization through Eglin AFB. The initial phase of the program is to build and test a development unit. The second phase is to integrate the startracker with the Charles Stark Draper Laboratory Micromechanical Inertial Guidance System (MIGS) and the Signal Processing Packaging Design (SPPD) being developed by Texas Instruments. The preliminary design review was conducted in November 1991. Three-axes prototype telescope assemblies have been built and design evaluation tests initiated.

Miller, Rick↗

NASA NTRS1993

Emerging technologies in microguidance and control

Employing recent advances in microfabrication, the Charles Stark Draper Laboratory has developed inertial guidance instruments of very small size and low cost. Microfabrication employs the batch processing techniques of solid state electronics, such as photolithography, diffusion, and etching, to carve mechanical parts. Within a few years, microfabricated gyroscopes should perform in the 10 to 100 deg/h range. Microfabricated accelerometers have demonstrated performance in the 50 to 500 microgravity range. These instruments will result in not only the redesign of conventional military products, but also new applications that could not exist without small, inexpensive sensors and computing. Draper's microfabricated accelerometers and gyroscopes will be described and test results summarized. Associated electronics and control issues will also be addressed. Gimballed, vibrating gyroscopes and force rebalance accelerometers constructed from bulk silicon, polysilicon surface-machined tuning fork gyroscopes, and quartz resonant accelerometers and gyroscopes are examined. Draper is pursuing several types of devices for the following reasons: to address wide ranges of performance, to realize construction in a flat pack, and to lessen the risks associated with emerging technologies.

Weinberg, Marc S.↗

NASA NTRS1999

Overview of Photonic Materials for Application in Space Environments

Future space systems will he based on components evolving from the development and refinement of new and existing photonic materials. Optically based sensors, inertial guidance, tracking systems, communications, diagnostics, imaging and high speed optical processing are but a few of the applications expected to widely utilize photonic materials. The response of these materials to space environment effects (SEE) such as spacecraft charging, orbital debris, atomic oxygen, ultraviolet irradiation, temperature and ionizing radiation will be paramount to ensuring successful space applications. The intent of this paper is to, address the latter two environments via a succinct comparison of the known sensitivities of selected photonic materials to the temperature and ionizing radiation conditions found in space and enhanced space environments Delineation of the known temperature and radiation induced responses in LiNbO3, AlGaN, AlGsAs,TeO2, Si:Ge, and several organic polymers are presented. Photonic materials are realizing rapid transition into applications for many proposed space components and systems including: optical interconnects, optical gyros, waveguide and spatial light modulators, light emitting diodes, lasers, optical fibers and fiber optic amplifiers. Changes to material parameters such as electrooptic coefficients, absorption coefficients, polarization, conductivity, coupling coefficients, diffraction efficiencies, and other pertinent material properties examined for thermo-optic and radiation induced effect. Conclusions and recommendations provide the reader with an understanding of the limitations or attributes of material choices for specific applications.

Taylor, E. W.↗

NASA NTRS1960

Orbital Capabilities of a Jupiter - Modified Agena "B" Space Vehicle

This presentation is directed towards providing additional information concerning the orbital capabilities for a two stage JUPITER-Agena "B" space vehicle. The data presented in this discussion is an extension of the preliminary information presented to NASA on 16 November, 1960. The pertinent features contained in this presentation are the usable payload and expected orbital accuracies associated with the vehicles under consideration when utilizing either an LEV-3 or ST-120 inertial guidance system. Missions of primary concern are 300, 400 or 500 nautical mile orbits around the earth. Preliminary budgetary and planning cost estimates related to the vehicle configurations discussed are presented under separate cover.

Chrysler Corporation↗