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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Linear Actuator Has Long Stroke and High Resolution

The term precision linear actuator, direct drive ( PLADD ) refers to a robust linear actuator designed to be capable of repeatedly performing, over a lifetime of the order of 5 to 10 years, positioning maneuvers that include, variously, submicron increments or slews of the order of a centimeter. The PLADD is capable of both long stroke (120 mm) and high resolution (repeatable increments of 20 nm). Unlike precise linear actuators of prior design, the PLADD contains no gears, levers, or hydraulic converters. The PLADD, now at the prototype stage of development, is intended for original use as a coarse-positioning actuator in a spaceborne interferometer. The PLADD could also be adapted to terrestrial applications in which there are requirements for long stroke and high resolution: potential applications include medical imaging and fabrication of semiconductor devices. The PLADD (see figure) includes a commercially available ball-screw actuator driven directly by a commercially available three-phase brushless DC motor. The ball-screw actuator comprises a spring-preloaded ball nut on a ball screw that is restrained against rotation as described below. The motor is coupled directly (that is, without an intervening gear train) to a drive link that, in turn, is coupled to the ball nut. By eliminating the gear train, the direct-drive design eliminates the complexity, backlash, and potential for misalignment associated with a gear train. To prevent inadvertent movement, there is a brake that includes flexured levers compressed against the drive link by preload springs. This is a power-off brake: There are also piezoelectric stacks that can be activated to oppose the springs and push the levers away from the drive link. Hence, power must be applied to the piezoelectric stacks to release the drive link from braking. To help ensure long operational life, all of the mechanical drive components are immersed in an oil bath within hermetically sealed bellows. The outer end of the bellows holds the outer end of the ball screw, thereby preventing rotation of the ball screw. Positioning is controlled by an electronic control system that includes digital and analog subsystems that interact with the motor and brake and with two sensor/encoder units: a Hall-effect-sensor rotation encoder and a linear glass-scale encoder. This system implements a proportional + integral + derivative control algorithm that results in variation of voltage commands to each of the three pairs of windings of the brushless DC motor. In one of two alternative control modes, the voltages are applied to the windings in a trapezoidal commutation scheme on the basis of timing signals obtained from the Hall-effect sensors; this scheme yields relatively coarse positioning - 24 steps per motor revolution. The second control mode involves a sinusoidal commutation scheme in which the output of the linear glass-scale encoder is transposed to rotational increments to yield much finer position feedback - more than 400,000 steps per revolution.

Cook, Brant T.↗
NASA NTRS2014

Aeroacoustics of Space Vehicles

While for airplanes the subject of aeroacoustics is associated with community noise, for space vehicles it is associated with vibro-acoustics and structural dynamics. Surface pressure fluctuations encountered during launch and travel through lower part of the atmosphere create intense vibro-acoustics environment for the payload, electronics, navigational equipment, and a large number of subsystems. All of these components have to be designed and tested for flight-certification. This presentation will cover all three major sources encountered in manned and unmanned space vehicles: launch acoustics, ascent acoustics and abort acoustics. Launch pads employ elaborate acoustic suppression systems to mitigate the ignition pressure waves and rocket plume generated noise during the early part of the liftoff. Recently we have used large microphone arrays to identify the noise sources during liftoff and found that the standard model by Eldred and Jones (NASA SP-8072) to be grossly inadequate. As the vehicle speeds up and reaches transonic speed in relatively denser part of the atmosphere, various shock waves and flow separation events create unsteady pressure fluctuations that can lead to high vibration environment, and occasional coupling with the structural modes, which may lead to buffet. Examples of wind tunnel tests and computational simulations to optimize the outer mold line to quantify and reduce the surface pressure fluctuations will be presented. Finally, a manned space vehicle needs to be designed for crew safety during malfunctioning of the primary rocket vehicle. This brings the subject of acoustic environment during abort. For NASAs Multi-Purpose Crew Vehicle (MPCV), abort will be performed by lighting rocket motors atop the crew module. The severe aeroacoustics environments during various abort scenarios were measured for the first time by using hot helium to simulate rocket plumes in the Ames unitary plan wind tunnels. Various considerations used for the helium simulation and the final confirmation from a flight test will be presented.

launch vehicle↗
NASA NTRS2000

Ion Thruster Used to Propel the Deep Space 1 Spacecraft to Comet Encounters

The NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) Project provided a xenon ion propulsion system to the Deep Space 1 (DS1) spacecraft to validate the propulsion system as well as perform primary propulsion for asteroid and comet encounters. The On-Board Propulsion Branch of the NASA Glenn Research Center at Lewis Field developed engineering model versions of the 30-cm-diameter ion thruster and the 2.5-kW power processor unit (PPU). Glenn then transferred the thruster and PPU technologies to Hughes Electron Dynamics and managed the contract, which supplied two flight sets of thrusters and PPU s to the Deep Space 1 spacecraft and to a ground-based life verification test at the Jet Propulsion Laboratory (JPL). In addition to managing the DS1 spacecraft development, JPL was responsible for the NSTAR Project management, thruster life tests, the feed system, diagnostics, and propulsion subsystem integration. The ion propulsion development team included NASA Glenn, JPL, Hughes Electronics, Moog Inc., and Spectrum Astro Inc. The overall NSTAR subsystem dry mass, including thruster, PPU, controller, cables, and the xenon storage and feed system, is 48 kg. The mass of the xenon stored onboard DS1 was about 81 kg, and the spacecraft wet mass was approximately 500 kg.The DS1 spacecraft was launched on October 24, 1998, and on July 29, 1999, it flew within 16 miles of the small asteroid Braille (formerly 1992KD) at a relative speed of 35,000 mph. As of November 1999, the ion propulsion system had performed flawlessly for nearly 149 days of thrusting. NASA has approved an extension to the mission, which will allow DS1 to continue thrusting to encounters with two comets in 2001. The DS1 optical and plasma diagnostic instruments will be used to investigate the comet and space environments. The spacecraft is scheduled to fly past the dormant comet Wilson- Harrington in January 2001 and the very active comet Borrelly in September 2001, at which time approximately 500 days of ion engine thrusting will have been completed.

Sovey, James S.↗
NASA NTRS1993

Solar Anomalous and Magnetospheric Particle Explorer attitude control electronics box design and performance

The attitude Control Electronics (ACE) Box is the center of the Attitude Control Subsystem (ACS) for the Solar Anomalous and Magnetospheric Particle Explorer (SAMPEX) satellite. This unit is the single point interface for all of the Attitude Control Subsystem (ACS) related sensors and actuators. Commands and telemetry between the SAMPEX flight computer and the ACE Box are routed via a MIL-STD-1773 bus interface, through the use of an 80C85 processor. The ACE Box consists of the flowing electronic elements: power supply, momentum wheel driver, electromagnet driver, coarse sun sensor interface, digital sun sensor interface, magnetometer interface, and satellite computer interface. In addition, the ACE Box also contains an independent Safehold electronics package capable of keeping the satellite pitch axis pointing towards the sun. The ACE Box has dimensions of 24 x 31 x 8 cm, a mass of 4.3 kg, and an average power consumption of 10.5 W. This set of electronics was completely designed, developed, integrated, and tested by personnel at NASA GSFC. SAMPEX was launched on July 3, 1992, and the initial attitude acquisition was successfully accomplished via the analog Safehold electronics in the ACE Box. This acquisition scenario removed the excess body rates via magnetic control and precessed the satellite pitch axis to within 10 deg of the sun line. The performance of the SAMPEX ACS in general and the ACE Box in particular has been quite satisfactory.

Chamberlin, K.↗
NASA NTRS2004

Detectors for the James Webb Space Telescope Near-Infrared Spectrograph

The Near-Infrared Spectrograph (NIRSpec) is the James Webb Space Telescope's primary near-infrared spectrograph. NASA is providing the NIRSpec detector subsystem, which consists of the focal plane array, focal plane electronics, cable harnesses, and software. The focal plane array comprises two closely-butted lambda (sub co) approximately 5 micrometer Rockwell HAWAII- 2RG sensor chip assemblies. After briefly describing the NIRSpec instrument, we summarize some of the driving requirements for the detector subsystem, discuss the baseline architecture (and alternatives), and presents some recent detector test results including a description of a newly identified noise component that we have found in some archival JWST test data. We dub this new noise component, which appears to be similar to classical two-state popcorn noise in many aspects, "popcorn mesa noise." We close with the current status of the detector subsystem development effort.

Rauscher, Bernard J.↗
NASA NTRS2008

Optical Coating Performance and Thermal Structure Design for Heat Reflectors of JWST Electronic Control Unit

The James Webb Space Telescope (JWST) consists of an infrared-optimized Optical Telescope Element (OTE) that is cooled down to 40 degrees Kelvin. A second adjacent component to the OTE is the Integrated Science Instrument Module, or ISIM. This module includes the electronic compartment, which provides the mounting surfaces and ambient thermally controlled environment for the instrument control electronics. Dissipating the 200 watts generated from the ISIM structure away from the OTE is of paramount importance so that the spacecraft's own heat does not interfere with the infrared light detected from distant cosmic sources. This technical challenge is overcome by a thermal subsystem unit that provides passive cooling to the ISIM control electronics. The proposed design of this thermal radiator consists of a lightweight structure made out of composite materials and low-emittance metal coatings. In this paper, we will present characterizations of the coating emittance, bidirectional reflectance, and mechanical structure design that will affect the performance of this passive cooling system.

Quijada, Manuel A.↗
NASA NTRS1993

Radiation and temperature effects on electronic components investigated under the CSTI high capacity power project

The effects of nuclear radiation and high temperature environments must be fully known and understood for the electronic components and materials used in both the Power Conditioning and Control subsystem and the reactor Instrumentation and Control subsystem of future high capacity nuclear space power systems. This knowledge is required by the designer of these subsystems in order to develop highly reliable, long-life power systems for future NASA missions. A review and summary of the experimental results obtained for the electronic components and materials investigated under the power management element of the Civilian Space Technology Initiative (CSTI) high capacity power project are presented: (1) neutron, gamma ray, and temperature effects on power semiconductor switches, (2) temperature and frequency effects on soft magnetic materials; and (3) temperature effects on rare earth permanent magnets.

Schwarze, Gene E.↗
NASA NTRS1979

Design package for programmable controller and hydronic subsystem

Information used in the evaluation of design of Sunkeeper Control's electronic controllers and hydronic packages is discussed. This information includes system performance specifications, a design data brochure, drawings, and qualification and acceptance test procedures.

Source record↗
NASA NTRS1979

Voyager 1 - Energetic ions and electrons in the Jovian magnetosphere

The observations of the cosmic-ray subsystem have added significantly to our knowledge of Jupiter's magnetosphere. The most surprising result is the existence of energetic sulfur, sodium, and oxygen nuclei with energies above 7 MeV per nucleon which were found inside of Io's orbit. Also, significant fluxes of similarly energetic ions reflecting solar cosmic-ray composition were observed throughout the magnetosphere beyond 11 times the radius of Jupiter. It was also found that energetic protons are enhanced by 30 to 70% in the active hemisphere. Finally, the first observations were made of the magnetospheric tail in the dawn direction out to 160 Jupiter radii.

Vogt, R. E.↗
NASA NTRS2002

Recent photonics activites under the NASA electronic parts and packaging (NEPP) program

With the rapidly increasing insertion of photonic devices, circuits and subsystems into NASA spacecraft, a variety of issues associated with reliability and radiation tolerance have arisen. In this paper, we discuss these issues from the perspective of the work currently ongoing in the NASA Electronic Parts and Packaging (NEPP) Program. This Program is focused on evaluating the reliability and radiation response of advanced and emerging microelectronics and photonics technologies of interest to NASA spacecraft system designers.

NEEP↗
NASA NTRS2002

The Microwave Anisotropy Probe (MAP) Guidance, Navigation, and Control Hardware Suite

The on-orbit success of the Microwave Anisotropy Probe (MAP) Guidance, Navigation, and Control System can partially be attributed to the performance of a hardware suite chosen to meet the complex attitude determination and control requirements of the mission. To meet these requirements, a diverse set of components was used. The set included two Lockheed Martin AST-201 star trackers, two Kearfott Two-Axis Rate Assemblies mounted to provide X, Y and redundant Z-axis rates, two Adcole Digital Sun Sensor heads sharing one set of electronics, twelve Adcole Coarse Sun Sensor eyes, three Ithaco E-sized Reaction Wheel Assemblies, a Propulsion Subsystem that employed eight Primex Rocket Engine Modules, and a pair of Goddard-designed Attitude Control Electronics which connect all of the components to the spacecraft processor. The performance of this hardware is documented, as are some of the spacecraft accommodations and lessons learned that came from working with this particular set of hardware.

Ward, David K.↗
NASA NTRS2015

BioSentinel: Mission Development of a Radiation Biosensor to Gauge DNA Damage and Repair Beyond Low Earth Orbit on a 6U Nanosatellite.

We are designing and developing a "6U" (10 x 22 x 34 cm; 14 kg) nanosatellite as a secondary payload to fly aboard NASA's Space Launch System (SLS) Exploration Mission (EM) 1, scheduled for launch in late 2017. For the first time in over forty years, direct experimental data from biological studies beyond low Earth orbit (LEO) will be obtained during BioSentinel's 12- to 18- month mission. BioSentinel will measure the damage and repair of DNA in a biological organism and allow us to compare that to information from onboard physical radiation sensors. In order to understand the relative contributions of the space environment's two dominant biological perturbations, reduced gravity and ionizing radiation, results from deep space will be directly compared to data obtained in LEO (on ISS) and on Earth. These data points will be available for validation of existing biological radiation damage and repair models, and for extrapolation to humans, to assist in mitigating risks during future long-term exploration missions beyond LEO. The BioSentinel Payload occupies 4U of the spacecraft and will utilize the monocellular eukaryotic organism Saccharomyces cerevisiae (yeast) to report DNA double-strand-break (DSB) events that result from ambient space radiation. DSB repair exhibits striking conservation of repair proteins from yeast to humans. Yeast was selected because of 1) its similarity to cells in higher organisms, 2) the well-established history of strains engineered to measure DSB repair, 3) its spaceflight heritage, and 4) the wealth of available ground and flight reference data. The S. cerevisiae flight strain will include engineered genetic defects to prevent growth and division until a radiation-induced DSB activates the yeast's DNA repair mechanisms. The triggered culture growth and metabolic activity directly indicate a DSB and its successful repair. The yeast will be carried in the dry state within the 1-atm P/L container in 18 separate fluidics cards with each card having 16 independent culture microwells, with integral microchannels and filters to supply nutrients and reagents, confine the yeast to the wells, and enable optical measurement. The measurement subsystem will monitor each subgroup of culture wells continuously for several weeks, optically tracking DSBtriggered cell growth and metabolism. BioSentinel will also include physical radiation sensors based on the TimePix sensor, as implemented by JSC's RadWorks group, which record individual radiation events including estimates of their linear-energytransfer (LET) values. Radiation-dose and LET data will be compared directly to the rate of DSB-and-repair events measured by the S. cerevisiae biosentinels. The spacecraft bus will operate in a deep space environment with functions that include command and data handling, communications, power generation (via deployable solar panels) and storage, and attitude determination-and-control system with micropropulsion. Development of the BioSentinel spacecraft will mature and prove multiple nanosatellite advances in order to function well beyond LEO: Communications from distances of ≥ 500,000 km; Autonomous attitude control, momentum management, and safe mode of nanosatellites in deep space; Shielding-, hardening-, design-, and software-derived radiation tolerance for electronics; Reliable functionality for 12 - 18 months of key subsystems for biofluidics, memory, communications, power, etc.; Close integration of living biological radiation event monitors with miniature physical radiation spectrometers; Biological measurement of solar particle events beyond Earth orbit In addition to providing the first biological results from beyond LEO in over 4 decades, BioSentinel will provide an adaptable small-satellite instrument platform to perform a range of human-exploration-relevant measurements that characterize the biological consequences of multiple outer space environments. BioSentinel is being developed under NASA's Advanced Exploration Systems program.

DNA damage↗
NASA NTRS2014

The MAVEN Magnetic Field Investigation

The MAVEN magnetic field investigation is part of a comprehensive particles and fields subsystem that will measure the magnetic and electric fields and plasma environment of Mars and its interaction with the solar wind. The magnetic field instrumentation consists of two independent tri-axial fluxgate magnetometer sensors, remotely mounted at the outer extremity of the two solar arrays on small extensions ("boomlets"). The sensors are controlled by independent and functionally identical electronics assemblies that are integrated within the particles and fields subsystem and draw their power from redundant power supplies within that system. Each magnetometer measures the ambient vector magnetic field over a wide dynamic range (to 65,536 nT per axis) with a quantization uncertainty of 0.008 nT in the most sensitive dynamic range and an accuracy of better than 0.05%. Both magnetometers sample the ambient magnetic field at an intrinsic sample rate of 32 vector samples per second. Telemetry is transferred from each magnetometer to the particles and fields package once per second and subsequently passed to the spacecraft after some reformatting. The magnetic field data volume may be reduced by averaging and decimation, when necessary to meet telemetry allocations, and application of data compression, utilizing a lossless 8-bit differencing scheme. The MAVEN magnetic field experiment may be reconfigured in flight to meet unanticipated needs and is fully hardware redundant. A spacecraft magnetic control program was implemented to provide a magnetically clean environment for the magnetic sensors and the MAVEN mission plan provides for occasional spacecraft maneuvers - multiple rotations about the spacecraft x and z axes - to characterize spacecraft fields and/or instrument offsets in flight.

Magnetic↗
NASA NTRS1977

The 26-meter S-X Conversion Project

The 26-meter S-X conversion project provides for the conversion of an existing 26-meter S-band subnet to a 34-meter S- and X-band subnet. The subnet chosen for conversion consists of the following stations: DSS 12 near Barstow, DSS 44 in Australia, and DSS 62 in Spain. The main subsystems effected by this project are the antenna mechanical, antenna microwave, and receiver-exciter. In addition to these, there are many project-related electronic equipments that have been added to the existing station equipment. The major subsystems are essentially through the design stage with the antenna mechanical subsystem completed through detail design with procurement in process.

Lobb, V. B.↗
NASA NTRS1975

Instruction manual, optical effects module electronic controller and processor, model OEMCP

The OEM-1 electronic module is discussed; it is comprised of four subsystems: the signal processing and display; the stepper motor controls; the chopper controls; and the dc-dc invertor. The OEM-1 module controls the sample wheel so that the relative transmittance of the samples can be compared to the clear aperture position. The 3-1/2 digit digital voltmeter displays the clear aperture signal level as well as the ratio of the remaining sample positions relative to the clear aperture position. The sample wheel position is decoded so that the signals and ratios can be correlated to the data. The OEM is automatically reset to the I sub o on initial turn-on and can be reset to the '0' position by actuating a front panel switch. The sample wheel can be interrupted to change samples or induce a longer integration time if desired by a front panel command. Integration times from 1 - 50 seconds are provided at the front panel, and BCD data for external interfacing is provided.

Source record↗
NASA NTRS1979

A thermal control approach for a solar electric propulsion thrust subsystem

A thrust subsystem thermal control design is defined for a Solar Electric Propulsion System (SEPS) proposed for the comet Halley Flyby/comet Tempel 2 rendezvous mission. A 114 node analytic model, developed and coded on the systems improved numerical differencing analyzer program, was employed. A description of the resulting thrust subsystem thermal design is presented as well as a description of the analytic model and comparisons of the predicted temperature profiles for various SEPS thermal configurations that were generated using this model. It was concluded that: (1) a BIMOD engine system thermal design can be autonomous; (2) an independent thrust subsystem thermal design is feasible; (3) the interface module electronics temperatures can be controlled by a passive radiator and supplementary heaters; (4) maintaining heat pipes above the freezing point would require an additional 322 watts of supplementary heating power for the situation where no thrusters are operating; (5) insulation is required around the power processors, and between the interface module and the avionics module, as well as in those areas which may be subjected to solar heating; and (6) insulation behind the heat pipe radiators is not necessary.

Maloy, J. E.↗
NASA NTRS1991

Rendezvous radar for the orbital maneuvering vehicle

The Rendezvous Radar Set (RRS) was designed at Motorola's Strategic Electronics Division in Chandler, Arizona, to be a key subsystem aboard NASA's Orbital Maneuvering Vehicle (OMV). The unmanned OMV, which was under development at TRW's Federal Systems Division in Redondo Beach, California, was designed to supplement the Shuttle's satellite delivery, retrieval, and maneuvering activities. The RRS was to be used to locate and then provide the OMV with vectoring information to the target satellite (or Shuttle or Space Station) to aid the OMV in making a minimum fuel consumption approach and rendezvous. The OMV development program was halted by NASA in 1990 just as parts were being ordered for the RRS engineering model. The paper presented describes the RRS design and then discusses new technologies, either under development or planned for development at Motorola, that can be applied to radar or alternative sensor solutions for the Automated Rendezvous and Capture problem.

Locke, John W.↗