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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 127 records · Page 7
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NASA NTRS1967

Manned Venus Orbiting Mission

Manned orbiting stopover round trips to Venus are studied for departure dates between 1975 and 1986 over a range of trip times and stay times. The use of highly elliptic parking orbits at Venus leads to low initial weights in Earth orbit compared with circular orbits. For the elliptic parking orbit, the effect of constraints on the low altitude observation time on the initial weight is shown. The mission can be accomplished with the Apollo level of chemical propulsion, but advanced chemical or nuclear propulsion can give large weight reductions. The Venus orbiting mission weights than the corresponding Mars mission.

LAUNCH DATE↗
NASA NTRS1974

Variable Side-Look Angle Concept For Radar Mapping

Radar mapping of planets can be accomplished at lower cost and with reduced emphasis on propulsion system capability from spacecraft operating in elliptical orbit than from circular orbit.

Vandenberg, Frank A.↗
NASA NTRS1978

System implementation for Earth Radiation Budget Satellite System

A description is presented of the instrument system which is needed for the Earth Radiation Budget Satellite System (ERBSS). The system is to be composed of instruments on two of NOAA's near-polar sun-synchronous Tiros-N/NOAA A through G series of operational satellites and on a NASA midinclination satellite of the Applications Explorer Mission (AEM) type referred to as ERBS-A/AEM. The Tiros-N/NOAA satellites will be in nominal 833 km altitude circular orbits with orbital inclinations of 98 deg. The AEM satellite will be in a circular orbit with an inclination of approximately 56 deg and a nominal altitude of 600 km. Each satellite will carry wide field-of-view (WFOV) and medium field-of-view (MFOV) sensors, a sensor for measuring the solar constant, and a narrow field-of-view (NFOV) cross-track scanner. The conceptual design of the W/MFOV instrument is discussed along with the conceptual design of the scanner.

Cooper, J. E.↗
NASA NTRS1989

Small Mercury Relativity Orbiter

The accuracy of solar system tests of gravitational theory could be very much improved by range and Doppler measurements to a Small Mercury Relativity Orbiter. A nearly circular orbit at roughly 2400 km altitude is assumed in order to minimize problems with orbit determination and thermal radiation from the surface. The spacecraft is spin-stabilized and has a 30 cm diameter de-spun antenna. With K-band and X-band ranging systems using a 50 MHz offset sidetone at K-band, a range accuracy of 3 cm appears to be realistically achievable. The estimated spacecraft mass is 50 kg. A consider-covariance analysis was performed to determine how well the Earth-Mercury distance as a function of time could be determined with such a Relativity Orbiter. The minimum data set is assumed to be 40 independent 8-hour arcs of tracking data at selected times during a two year period. The gravity field of Mercury up through degree and order 10 is solved for, along with the initial conditions for each arc and the Earth-Mercury distance at the center of each arc. The considered parameters include the gravity field parameters of degree 11 and 12 plus the tracking station coordinates, the tropospheric delay, and two parameters in a crude radiation pressure model. The conclusion is that the Earth-Mercury distance can be determined to 6 cm accuracy or better. From a modified worst-case analysis, this would lead to roughly 2 orders of magnitude improvement in the knowledge of the precession of perihelion, the relativistic time delay, and the possible change in the gravitational constant with time.

Bender, Peter L.↗
NASA NTRS1991

Shuttle-C auxiliary propulsion system certification approach and low-cost evolution concepts

Shuttle-C provides NASA with a heavy-lift cargo capability that is a low-cost derivative of the current Space Shuttle. This system can deliver 80,000 to 140,000 pounds of payload to low earth orbit within 4 years of authority to proceed. Shuttle-C will share th existing Shuttle launch facilities at Kennedy Space Center (KSC). The Shuttle-C vehicle incorporates an auxiliary propulsion system (APS) which provides separation from the external tank (ET), orbit circularization, on-orbit attitude hold, and deorbit capability. The APS is derived from the Shuttle orbiter aft reaction control system (ARCS) and orbital maneuvering system (OMS) with the objective of achieving minimum life cycle cost for this expendable vehicle. The resulting APS system will be described, its plan for a minimum-cost flight certification utilizing a cold-flow test program explained, and low-cost evolution by component design modification and part substitution presented.

Harsh, M. George↗
NASA NTRS1991

Hubble Space Telescope (HST): Emergency support

The Hubble Space Telescope (HST) is a national facility that consists of a 2.4-m aperture Ritchey-Chretien cassegrain telescope weighing approximately 9525 kg with various energy detectors designed for the observation of IR, visible, and UV wavelengths (0.12 to 1000 microns). The HST was deployed into a 28.5-degree inclination, circular orbit, which permits a mission lifetime of 15 years. The orbit is decaying circular between 594 and 400 km x 28.5 deg; the period = 95 minutes. Information is presented in tabular form on the following areas: Deep Space Network support, frequency assignments, telemetry, command, and tracking support responsibility.

Repass, J.↗
NASA NTRS1998

Martian Meteorology: Determination of Large Scale Weather Patterns from Surface Measurements

We employed numerical modelling of the martian atmosphere, and our expertise in understanding martian atmospheric processes, to better understand the coupling between lower and upper atmosphere processes. One practical application of this work has been our involvement with the ongoing atmospheric aerobraking which the Mars Global Surveyor (MGS) spacecraft is currently undergoing at Mars. Dr. Murphy is currently a member of the Mars Global Surveyor (MGS) Aerobraking Atmospheric Advisory Group (AAG). He was asked to participate in this activity based upon his knowledge of martian atmospheric dynamical processes. Aerobraking is a process whereby a spacecraft, in an elliptical orbit, passes through the upper layers of the atmosphere (in this instance Mars). This passage through the atmosphere 'drags' upon the spacecraft, gradually reducing its orbital velocity. This has the effect, over time, of converting the elliptical orbit to a circular orbit, which is the desired mapping orbit for MGS. Carrying out aerobraking eliminates the need for carrying large amounts of fuel on the spacecraft to execute an engine burn to achieve the desired orbit. Eliminating the mass of the fuel reduces the cost of launch. Damage to one of MGS's solar panels shortly after launch has resulted in a less aggressive extended in time aerobraking phase which will not end until March, 1999. Phase 1 extended from Sept. 1997 through March 1998. During this time period, Dr. Murphy participated almost daily in the AAG meetings, and beginning in December 1997 lead the meeting several times per week. The leader of each of the daily AAG meetings took the results of that meeting current state of the atmosphere, identification of any time trends or spatial patterns in upper atmosphere densities, etc.) forward to the Aerobraking Planning Group (APG) meeting, at which time the decision was made to not chance MGS orbit, to lower the orbit to reach higher densities (greater 'drag'), or raise the orbit to avoid experiencing excessive, possibly damaging densities.

Murphy, James R.↗
NASA NTRS1998

Martian Meteorology: Determination of Large Scale Weather Patterns from Surface Measurements

We employed numerical modelling of the martian atmosphere, and our expertise in understanding martian atmospheric processes, to better understand the coupling between lower and upper atmosphere processes. One practical application of this work has been our involvement with the ongoing atmospheric aerobraking which the Mars Global Surveyor (MGS) spacecraft is currently undergoing at Mars. Dr. Murphy is currently a member of the Mars Global Surveyor (MGS) Aerobraking Atmospheric Advisory Group (AAG). He was asked to participate in this activity based upon his knowledge of martian atmospheric dynamical processes. Aerobraking is a process whereby a spacecraft, in an elliptical orbit, passes through the upper layers of the atmosphere (in this instance Mars). This passage through the atmosphere 'drags'upon the spacecraft, gradually reducing its orbital velocity. This has the effect, over time, of converting the elliptical orbit to a circular orbit, which is the desired mapping orbit for MGS. Carrying out aerobraking eliminates the need for carrying large amounts of fuel on the spacecraft to execute an engine burn to achieve the desired orbit. Eliminating the mass of the fuel reduces the cost of launch. Damage to one of MGS's solar panels shortly after launch has resulted in a less aggressive extended in time aerobraking phase which will not end until March, 1999. Phase I extended from Sept. 1997 through March 1998. During this time period, Dr. Murphy participated almost daily in the AAG meetings, and beginning in December 1997 lead the meeting several times per week. The leader of each of the daily AAG meetings took the results of that meeting (current state of the atmosphere, identification of any time trends or spatial patterns in upper atmosphere densities, etc.) forward to the Aerobraking Planning Group (APG) meeting, at which time the decision was made to not change MGS orbit, to lower the orbit to reach higher densities (greater 'drag'), or raise the orbit to avoid experiencing excessive, possibly damaging densities.

Murphy, James R.↗
NASA NTRS2018

In-Flight Operation of the Dawn Ion Propulsion System Through Completion of the Final Orbit Around Dwarf Planet Ceres

The Dawn mission, part of NASA’s Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional DV of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in August 2012 and completed in late December 2014. From December 2014 through June 2016 the IPS was used for transiting the spacecraft to all science orbits at Ceres including the final orbit for Dawn’s primary mission, called the low altitude mapping orbit (LAMO), a circular orbit at a mean altitude above Ceres of approximately 385 km. Dawn met or exceeded all pre-launch science requirements and Dawn's prime mission concluded on June 30, 2016. Dawn subsequently received NASA approval for two extended missions at Ceres, called XM1 and XM2. During XM1 IPS was used for orbit maintenance and to transit the spacecraft to several new, higher-altitude science orbits, ending in an elliptical orbit at a maximum altitude of approximately 38,000 km. In XM2, IPS operations included transiting the Dawn spacecraft to its final, highly elliptical orbit at Ceres ranging from 35 km perigee to 4,800 km apogee. Science data acquisition will continue in XM2 until the hydrazine is exhausted, which is expected to occur between August and October 2018. Dawn has successfully completed all science goals for both the primary and XM1 extended missions. To date the IPS has been operated for approximately 51,250 hours, consumed approximately 416 kg of xenon, and provided a delta-V of almost 11.5 km/s, a record for an on-board propulsion system. The IPS performance characteristics are close to the expected performance based on analysis and testing performed pre-launch. Dawn’s IPS continues to be fully operational as of June 2018. This paper provides an overview of Dawn’s mission objectives and the results of Dawn IPS mission operations for XM1, and XM2 through June 2018.

Rayman, Marc D.↗
NASA NTRS1988

User and technical documentation

The program LIBRATE calculates velocities for trajectories from low earth orbit (LEO) to four of the five libration points (L2, L3, L4, and L5), and from low lunar orbit (LLO) to libration points L1 and L2. The flight to be analyzed departs from a circular orbit of any altitude and inclination about the Earth or Moon and finishes in a circular orbit about the Earth at the desired libration point within a specified flight time. This program produces a matrix of the delta V's needed to complete the desired flight. The user specifies the departure orbit, and the maximum flight time. A matrix is then developed with 10 inclinations, ranging from 0 to 90 degrees, forming the columns, and 19 possible flight times, ranging from the flight time (input) to 36 hours less than the input value, in decrements of 2 hours, forming the rows. This matrix is presented in three different reports including the total delta V's, and both of the delta V components discussed. The input required from the user to define the flight is discussed. The contents of the three reports that are produced as outputs are also described. The instructions are also included which are needed to execute the program.

Source record↗
NASA NTRS1999

Investigating the Martian Environment with the Mars Global Surveyor Thermal Emission Spectrometer

The Thermal Emission Spectrometer (TES) onboard Mars Global Surveyor (MGS) is being used to investigate the surface and atmosphere of Mars and the martian moons Phobos and Deimos. As such, it builds upon infrared observations obtained by previous spacecraft, e.g.. Mariner 9 and Viking Orbiters. The objectives of the TES experiment are: (1) determine and map surface minerals. rocks. and ices; (2) study the atmospheric dust composition. particle size, and spatial and temporal distribution; (3) investigate condensate clouds, CO2 and H2O, location. temperature, and height; (4) investigate polar cap deposits, e.g., growth, retreat, and energy balance; (5) measure the thermo-physical properties of surface materials; and (6) characterize the atmospheric structure and dynamics. The TES instrument is based upon a Michelson interferometer and collects data in the 1700-200 per cm region (about 6-50 micron at 5 or 10 per cm resolution. There are also broad-band bolometric (4.5100 microns and solar reflectance (0.3-2.7 microns) channels. The TES was designed to have a noise equivalent spectral radiance of 1.2 x l0(exp -8) per W per square cm per sr per cm corresponding to a signal-to-noise ratio of 490 at 1000 per cm (10 mm) for a 270 K scene and preflight data suggest a radiometric accuracy of about 1.2 x 10(exp -8) per W per square cm per sr per cm. In-flight observations indicate a small systematic calibration offset of about 1.2 x 10(exp -7) per W per square cm per sr per cm is present in the TES data. MGS achieved Mars orbital insertion September 11, 1997, and entered the initial aerobraking phase. MGS should have reached a circular orbit by early 1998. However, structural damage to one solar panel occurred during its deployment in the cruise phase and ultimately required an assessment of the extent of the damage and a much slower aerobraking period. This has delayed reaching the final circular orbit until March 1999. TES and the other MGS science instruments began operating just after orbital insertion and continued to obtain data until November 1998 when the TES was turned off to reduce power consumption on the spacecraft. TES data obtained during orbits 2 to 53 have previously been published and clearly illustrate the variety of scientific questions that can be addressed with them.

Roush, T. L.↗
NASA NTRS1976

Effects of solar radiation on the orbits of small particles

A modification of the Robertson (1937) equations of particle motion in the presence of solar radiation is developed which allows for partial reflection of sunlight as a result of rapid and varying particle rotations caused by interaction with the solar wind. The coefficients and forces in earlier forms of the equations are compared with those in the present equations, and secular rates of change of particle orbital elements are determined. Orbital dimensions are calculated in terms of time, probable sizes and densities of meteoric and cometary particles are estimated, and times of infall to the sun are computed for a particle moving in an almost circular orbit and a particle moving in an elliptical orbit of high eccentricity. Changes in orbital elements are also determined for particles from a long-period sun-grazing comet. The results show that the time of infall to the sun from a highly eccentric orbit is substantially shorter than from a circular orbit with a radius equal to the mean distance in the eccentric orbit. The possibility is considered that the free orbital kinetic energy of particles drawn into the sun may be the energy source for the solar corona.

Lyttleton, R. A.↗
NASA NTRS1994

Requirements report for SSTO vertical take-off/horizontal landing vehicle

This document describes the detailed design requirements and design criteria to support Structures/TPS Technology development for SSTO winged vehicle configurations that use vertical take-off and horizontal landing and deliver 25,000 lb payloads to a 220 nm circular orbit at an inclination of 51.6 degrees or 40,000 lb payloads to a 150 nm circular orbit at a 28.5 degree of inclination. This document will be updated on a timely basis as informatIon becomes available throughout the project.

Greenberg, H. S.↗
NASA NTRS1994

Requirements report for SSTO vertical take-off and horizontal landing vehicle

This document describes the detailed design requirements and design criteria to support Structures/TPS Technology development for SSTO winged vehicle configurations that use vertical take-off and horizontal landing and delivers 25,000 lb payloads to a 220 nm circular orbit at an inclination of 51.6 degrees or 40,000 lb payloads to a 150 nm circular orbit at a 28.5 degree inclination.

Greenberg, H. S.↗
NASA NTRS1988

PLANECHG: Earth orbit to lunar orbit delta V estimation program. User and technical documentation

The PLANECNG computer program calculates velocities for Earth-to-Mooon and Moon-to-Earth trajectories. The flight to be analyzed originates in a circular orbit of any inclination and altitude about one of the bodies, and culminates in a circular orbit of any inclination and altitude about the other body. An intermedate delta V and plane change occurs at the Lunar Sphere of Influence (SOI), the region where the vehicle is near its lowest velocity in the trajectory, and therefore where it is able to make the plane change with the lowest delta V. A given flight may penetrate the SOI at a number of points. Each point has associated with it a unique set of delta V's and total velocity. The program displays the velocities, in matrix form, for a representative set of SOI penetration points. An SOI point is identified by projecting Lunar latitude and longitude onto the SOI. The points recorded for a given flight are defined by the user, who provides a starting longitude and latitude, and an increment for each. A matrix is built with 10 longitudes forming the columns and 19 latitudes forming the rows. This matrix is presented in six reports, each containing different velocity or node information in the body of the matrix.

Source record↗
NASA NTRS1976

Dispersion of ensembles of non-interacting particles

The dynamics of an ensemble of noninteracting particles dispersing from a common origin and moving in a common force field with an initial distribution of momenta is analyzed using an approach where the particles are considered as a continuum described by a phase-space distribution function. General solutions are obtained for both the distribution function and the associated spatial density function. The linear case of small departures from circular orbits in an axisymmetric gravitational field is treated along with the specific case of particle dispersion from an object in a circular orbit in the same type of field. Numerical results are presented for the latter case, and consideration is given to the inverse problem of determining the initial time and velocity distribution from knowledge of the ensemble structure at a later time. Explicit results are provided for the case of an ellipsoidal distribution of initial momenta, and a numerical procedure is indicated for treating more general cases.

Heard, W. B.↗
NASA NTRS1985

Origin of mesosiderites as a natural consequence of planet formation

The mineral composition of mesosiderites is described and a theory of the origin and evolution of these meteorites is presented. It is suggested that the asteroid parent body of the mesosiderites also formed in the inner solar system, perhaps just within the orbit of Mars. As a result of close planetary encounters, some bodies that formed near Earth or Venus were gravitationally perturbed into non-circular orbits; a few such bodies passed through the mesosiderite region at high relative velocities, colliding with and destroying a few of the native asteroids. Olivine-rich silicate mantles shattered into small pieces, but the stronger metal cores remained as large fragments. Much of the debris remained in circular orbits and accreted to the basaltic regoliths of intact native asteroids at low relative velocities. The large core fragments that collided with the crust greatly enriched restricted regions of the surface in metal. These localized regions were the mesosiderite progenitors; they accounted for only about 1% of the surface area of the parent bodies.

Wasson, J. T.↗