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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

The effects of trajectory characteristics on scientific objectives for major planetary orbiters.

Outline of specific solutions in the areas of attitude control, interplanetary trajectory parameters, and orbital configurations of major planetary spacecraft. It is shown that inertia wheel supplementation to a three-axis, stabilized attitude control system will substantially reduce propellant weight and attitude jet firings, especially when considering three-year orbital lifetimes as well as interplanetary trip times of several years. Eight interplanetary trajectory parameters including the earth injection energy and the declination of the launch asymptote are shown for a 1980 earth launch mission to Jupiter. The effects of the interplanetary trajectory parameters on the deboost velocities and resulting orbits about Jupiter are summarized. Orbital mode and science interfaces are defined, and a celestial mechanics experiment independent of earth tracking is briefly outlined.

Paul, C. K.↗
NASA NTRS1992

Cosmic-ray exposure history of ordinary chondrites

The exposure age histograms for H-, L-, and LL-chondrites are discussed. None of these histograms is consistent with a continuous delivery of asteroidal material to the earth, as the observed T(e) histograms clearly disagree with expected exponential distributions for a variety of orbital lifetimes. It is concluded that T(e) histograms are dominated by stochastic events and that the continuous supply of asteroidal material can account only for a minor background of the T(e) histograms. An attempt is made to identify major collisional levels among the major classes of ordinary chondrites in order to estimate the frequency of stochastic events. Orbital maturity in the inner solar system is documented by the p.m./total fall ratio among observed meteorite falls. All chondrite classes exhibit a uniform ratio of 2/3 except type H5 chondrites, which reveal a 0.5 or lower p.m./total fall ratio. This shift in the time of fall statistics suggests a strongly evolved orbit for the H5 parent at the time of collision about 7 Ma ago.

Marti, K.↗
NASA NTRS2000

The Fast Alternative Cryogenic Experiment Testbed

One of the challenges in the area of cryogenics for space exploration in the next millennium is providing the capability for inexpensive, frequent, access to space. Faced with this challenge during the International Space Station (ISS) build era, when other Space Shuttle manifesting opportunities are unavailable, a "proof of concept" cryostat has been developed to demonstrate the ability to accommodate low temperature science investigations within the constraints of the Hitchhiker siderail carrier. The Hitchhiker siderail carrier is available on a "mass available" basis during the ISS build era. In fact, several hitchhiker payloads flew with the deployment of the Unity module. Hitchhiker siderail carrier payloads have historically flown an average of about four times a year. A hybrid Solid Neon - Superfluid Helium cryostat has been developed with Janis Research Company to accommodate instruments of 16.5 cm diameter and 30 cm. length. This hybrid approach was taken in part to provide adequate on-orbit lifetime for instruments with high (conducted) heat loads from the instrumentation wiring. Mass, volume, lifetime and the launch hold scenario were all design drivers. In addition, with Ball Aerospace and Technologies Corporation, a multichannel VME architecture Germanium Resistance Thermometer (GRT) readout and heater control servo system has been developed. In a flight system, the cryostat and electronics payloads would be umbilically attached in a paired Hitchhiker siderail mount, and permit on-orbit command and telemetry capability. The results of performance tests of both the cryostat, and a helium sample instrument will be presented. The instrument features a self contained, miniaturized, nano-Kelvin resolution High Resolution Thermometer (HRT). This high level of thermal resolution is achieved through the utilization of a dc Superconducting Quantum Interference Device (SQUID). Although developed for the Low Temperature Microgravity Fundamental Physics investigator community, many design features are applicable in fields such as infrared and x-ray astronomy.

Nash, Alfred↗
NASA NTRS1993

Lunar Scout Two spacecraft gravity experiment

Measurement of the gravity field of the Moon has a high science priority because of its implications for the internal structure and thermal history of the Moon, and it has a high priority for future exploration activities because of the influence of lunar gravity on spacecraft navigation and orbit maintenance. The current state of knowledge in the lunar gravity field (and the uncertainty in the knowledge) is based primarily on data accumulated from the Lunar Orbiter and Apollo programs. Data are sparse and emphasize the equatorial band (+/- 30 deg) on the near side of the Moon. There are no tracking data on the far side and only the Lunar Orbiter 5 provides a small amount of high inclination data. A host of gravity models developed from different combinations of tracking data have large discrepancies in their predictions of spacecraft motion and orbit lifetimes. There are also large disagreements in the Mercator projections of the gravity acceleration from each model, especially on the far side, where the contours tend to have no obvious relationship with the local topography. The science and engineering requirements for global gravity field mapping will be satisfied with continuous radio metric tracking of Lunar Scout 1 in a low polar orbit using the Deep Space Network and Lunar Scout 2 in a high elliptical orbit. The gravity field of the Moon will be mapped during the Scout Program using a two spacecraft concept. In the two spacecraft concept, one spacecraft is placed in a high altitude eccentric orbit while the second spacecraft is in a low altitude polar orbit. The gravity experiment requires a radio frequency that will permit two-way Doppler tracking between the spacecraft and the Deep Space Network (DSN). Both spacecraft carry NASA standard transponder systems for data transmission to Earth as well as for tracking and orbit determination. Data sufficient to produce a gravity field map could be acquired within one month with this system.

Cheng, Andrew F.↗
NASA NTRS2015

Cryogenic Orbital Testbed (CRYOTE) Ground Test Article, Final Report

Liquid propulsion has been used since Robert Goddard started developing a liquid oxygen (LO2) and gasoline powered rocket and fired it in 1923 (Ref. 1). In the following decades engineers settled on the combination of liquid hydrogen (LH2) and LO2 as the most efficient propellant combination for in-space travel. Due to their low temperatures (LH2 at 20 K and LO2 at 90 K), they require special handling and procedures. General Dynamics began developing LO2 and LH2 upper stages in 1956 in the form of Centaur, these efforts were soon funded by the Department of Defense in conjunction with NASA (beginning in 1958) (Ref. 2). Meanwhile NASA also worked with McDonnell Douglas to develop the SIV-B stage for the Saturn V rocket. In the subsequent years, the engineers were able to push the Centaur to up to 9 hr of orbital lifetime and the SIV-B to up to 6 hr. Due to venting the resultant boil-off from the high heat loads through the foam insulation on the upper stages, both vehicles remained in a settled configuration throughout the flights, thus the two phases of propellant (liquid and vapor) were separated at a known location. The one exception to this were the Titan/Centaur missions, which thanks to the lower boil-off using three layers of multilayer insulation (MLI), were able to coast unsettled for up to 5.25 hr during direct geosynchronous orbit insertion missions. In the years since there has been a continuous effort to extend the life of these upper stages from hours to days or even months.

Joule-Thompson Effect↗
NASA NTRS2018

Green Propulsion Advancement and Infusion

All space missions benefit from increased propulsion system performance, allowing lower spacecraft launch mass, larger scientific payloads, or extended on-orbit lifetimes. Likewise, long-term storable liquid propellant candidates that offer significant reduction in personnel hazards and shorter payload processing schedules present a more attractive propulsion subsystem solution to spacecraft builders. Aiming to reduce risk to potential infusion missions and fully comprehend the alternative propellant performance, the work presented herein represents many years of development and collaborative efforts to successfully align higher performance, low toxicity green propellants into NASA Goddard Space Flight Center (GSFC) missions. High Performance Green Propulsion (HPGP), and the associated propellant technology, has advanced significantly in maturity with increased familiarity with LMP-103S propellant handling, the proven reduction in loading hazards, successful launches conducted at multiple international Ranges, and HPGP on-orbit flight heritage. As science missions move forward to the potential infusion of HPGP technology, the National Aeronautics and Space Administration (NASA) and its partners are working to address gaps in system performance and operational considerations.

Spacecraft Propulsion and Power↗
NASA NTRS1992

The geocentric particulate distribution: Cometary, asteroidal, or space debris?

Definition of the Low Earth Orbit (LEO) particulate environment has been refined considerably with the analysis of data from NASA's Long Duration Exposure Facility (LDEF). Measurements of the impact rates from particulates ranging from sub-micron to millimetres in dimension and, especially, information on their directionality has permitted new scrunity of the sources of the particulates. Modelling of the dynamics of both bound (Earth orbital) and unbound (hyperbolic interplanetary) particulates intercepting LDEF's faces leads to the conclusion that the source is dominantly interplanetary for particle dimensions of greater than some 5 microns diameter; however the anisotropy below this dimension demands lower velocities and is compatible with an orbital component. Characteristics of the LDEF interplanetary component are compatible with familiar meteoroid sources and deep space measurements. Understanding of the orbital component which exceeds the interplanetary flux by a factor of 4 is less clear; although the very small particulates in orbit have been associated with space debris (Lawrance and Brownlee, 1986) this data conflicts with other measurements (McDonnell, Carey and Dixon, 1984) at the same epoch. By analysis of trajectories approaching the Earth and its atmosphere, we have shown that a significant contribution could be captured by aerocapture, i.e., atmospheric drag, from either asteroidal or cometary sources; such enhancement is unlikely however to provide the temporal and spatial fluctuations observed by the LDEF Interplanetary Dust Experiment (Mullholland et al. 1992). A further new mechanism is also examined, that of aerofragmentation capture, where an atmospheric grazing trajectory, which would not normally lead to capture, leads to fragmentation by thermal or mechanical shock; the microparticulates thus created can be injected in large numbers, but only into short-lifetime orbits. The concentration in one particular orbit plane, could explain the temporal fluctuations seen on LDEF; space debris could also explain the phenomenon.

Mcdonnell, J. A. M.↗
NASA NTRS2005

Mission Analysis for LEO Microwave Power-Beaming Station in Orbital Launch of Microwave Lightcraft

A detailed mission analysis study has been performed for a 1 km diameter, rechargeable satellite solar power station (SPS) designed to boost 20m diameter, 2400 kg Micr,oWave Lightcraft (MWLC) into low earth orbit (LEO) Positioned in a 476 km daily-repeating oi.bit, the 35 GHz microwave power station is configured like a spinning, thin-film bicycle wheel covered by 30% efficient sola cells on one side and billions of solid state microwave transmitter elements on the other, At the rim of this wheel are two superconducting magnets that can stor,e 2000 G.J of energy from the 320 MW, solar array over a period of several orbits. In preparation for launch, the entire station rotates to coarsely point at the Lightcraft, and then phases up using fine-pointing information sent from a beacon on-board the Lightcraft. Upon demand, the station transmits a 10 gigawatt microwave beam to lift the MWLC from the earth surface into LEO in a flight of several minutes duration. The mission analysis study was comprised of two parts: a) Power station assessment; and b) Analysis of MWLC dynamics during the ascent to orbit including the power-beaming relationships. The power station portion addressed eight critical issues: 1) Drag force vs. station orbital altitude; 2) Solar pressure force on the station; 3) Station orbital lifetime; 4) Feasibility of geo-magnetic re-boost; 5) Beta angle (i..e., sola1 alignment) and power station effective area relationship; 6) Power station percent time in sun vs, mission elapsed time; 7) Station beta angle vs.. charge time; 8) Stresses in station structures.. The launch dynamics portion examined four issues: 1) Ascent mission/trajecto1y profile; 2) MWLC/power-station mission geometry; 3) MWLC thrust angle vs. time; 4) Power station pitch rate during power beaming. Results indicate that approximately 0 58 N of drag force acts upon the station when rotated edge-on to project the minimum frontal area of 5000 sq m. An ion engine or perhaps an electrodynamic thruster (i.e., geomagnetic re-boost) station-keeping system can maintain the orbit altitude. The rate at which the power station s superconducting magnetic energy storage system (SMES) is 'charged' directly relates to the beta angle since the station is operating in the edge-on attitude. The maximum charge rate occurs when the beta angle is at its maximum because time in the sun and projected area of the station are, too, at their maximums For the maximum charge of 2000 G.J with a maximum beta angle of 52 degrees, approximately 3 hours (2 orbital revolutions) are required to reach the full charge, while about 16 hours (10.3 revolutions) are required when the beta angle is 10 degrees. Overall, the LEO station concept appears to be a viable candidate fo1 the formidable power-beaming infrastructure needed to boost MWLC into low earth orbit.

Myrabo, L. N.↗
NASA NTRS1971

The Smithsonian Earth Physics Satellite (SEPS) definition study, volumes 1 through 4

A limited Phase B study was undertaken to determine the merit and feasibility of launching a proposed earth physics satellite with Apollo-type hardware. The study revealed that it would be feasible to launch this satellite using a S-IB stage, a S-IVB with restart capability, an instrument unit, a SLA for the satellite shroud, and a nose cone (AS-204 configuration). A definition of the proposed satellite is provided, which is specifically designed to satisfy the fundamental requirement of providing an orbiting benchmark of maximum accuracy. The satellite is a completely passive, solid 3628-kg sphere of 38.1-cm radius and very high mass-to-area ratio (7980 kg sq mi). In the suggested orbit of 55 degrees inclination, 3720 km altitude, and low eccentricity, the orbital lifetime is extremely long, so many decades of operation can be expected.

Source record↗
NASA NTRS1973

Development of a rotating gravity gradiometer for earth orbit applications (AAFE)

Some preliminary mission studies are described along with the design, fabrication, and test of a breadboard model of an earth orbital, rotating gravity gradiometer with a design goal of 10 to the minus 11th power/sec sq (0.01 EU) in a 35-sec integration time. The proposed mission uses a Scout vehicle to launch one (or two orthogonally oriented) spin-stabilized satellites into a 330-km circular polar orbit some 20 days before an equinox. During the short orbital lifetime, the experiment would obtain two complete maps of the gravity gradient field with a resolution approaching 270 km (degree 75). The breadboard model of the gradiometer demonstrated a combined thermal and electronic noise threshold of 0.015 EU per data channel. The design changes needed to reduce the noise to less than 0.01 EU were identified. Variations of the sensor output signal with temperature were experimentally determined and a suitable method of temperature compensation was developed and tested. Other possible error sources, such as sensor interaction with satellite dynamics and magnetic fields, were studied analytically and shown to be small.

Forward, R. L.↗
NASA NTRS1974

A Ge/Li/ spectrometer for gamma ray astronomy

A space-borne instrument to measure the spectral and temporal characteristics of astronomical gamma rays in the .05 to 10 Mev energy range is described. The spectrometer consists of four 60 cc Ge(Li) detectors with an energy resolution of 2.5 Kev (FWHM) at 1.33 Mev. The Ge(Li) detectors are actively shielded with CsI(Na) for angular collimation and background reduction. Cooling of the Ge(Li) detectors to 88 K for orbital lifetimes of up to two years is provided by a two-stage solid subliming refrigerator.

Hicks, D. B.↗
NASA NTRS1975

The Atmosphere Explorer spacecraft system

The Atmosphere Explorer Spacecraft incorporates an orbiting aeronomy laboratory designed to operate in the lower thermosphere. A three-axis stabilized platform capable of operating over a wide range of orbital altitudes, inclinations, solar aspect angles, and spin rates satisfies the science requirements for instrument orientation. An on-board propulsion subsystem allows operations at low-perigee altitudes without severe limitation of orbital lifetime. Programmable high-speed digital command and data handling, together with power supplied by solar cells and batteries, allows the spacecraft to be operated remotely in daylight or darkness. This configuration of spacecraft support subsystems fulfills the requirements for the various phases of the scientific mission.

De Bastos, R.↗
NASA NTRS1975

Size distribution of fragment debris produced by simulated meteoroid impact of spacecraft wall

A laboratory experimental program was conducted to determine the amount of fragment debris produced and its size distribution when simulated spacecraft walls are penetrated by hypervelocity projectiles. Two tests were made in which a spacecraft wall was struck with a small steel and an aluminum projectile, respectively. Results show that most of the fragment debris were irregularly shaped flat plates. The fragment data follow a cumulative number distribution law of the form N = am(exp b) where m is fragment mass and a and b are constants. Orbital lifetimes of most of the fragments were very short, being on the order of a few days.

T Dale Bess↗