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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Seismoball: A Small Europa Orbiter Drop-Off Probe for Early Exploration of the Europan Surface

Recent magnetometry data received from Galileo indicate that the most likely explanation for the magnetic signature there is indeed a global conducting layer below the surface. This conducting layer is well matched by a salty, mineral rich strata beneath the Europan ice crust or a salt water ocean. Galileo imaging results show a variety of terrain types thought to contain young material; for example, lineaments, chaotic terrain, and eruption features. Additionally, Galileo images have shown indications of areas of up-welling where subsurface material periodically gets pushed to the surface due to the forces of fracturing, butting, and refreezing of the ice sheet. While Europa Orbiter will provide close-flyby high resolution images, as well as magnetometry, spectroscopy and other remote sensing data of the surface, it will not be able to provide essential engineering data like surface hardness and surface ice structure needed to support eventual landed missions. Additionally, ice chemical composition at microscopic scales can only be studied in detail through in situ instrumentation. Seismoball is a small probe designed to be injected into a surface intersect orbit around Europa. Using small reverse thrusters, the probe will be capable of nulling the high horizontal injection velocity as it approaches the 2 km surface injection altitude, thus allowing it to fall to the surface at an impact velocity of < 100m/sec (much less than the DS-2 impact velocities). The external breakaway thruster structure and crushable exterior shell absorb the impact energy while allowing the science instrument suite to remain intact. JPL has already started analyzing the entry dynamics and designing/building a small, low mass probe which will withstand the impact g-forces and fit as a 'carry-on' on board the Europa Orbiter. The probe will carry a suite of 5-6 micro-instruments for imaging the surface (both microscopic and far-field), surface and shallow subsurface ice temperatures, surface hardness, crustal dynamics and periodicity, and compositional chemistry. If selected, this flight development activity will provide a unique science opportunity and adjunct to the primary Orbiter science mission. The final flight system will be designed to accommodate orbiter mass, volume, and power interface constraints, as well as entry dynamics, g-load mitigation, and arbitrary landing orientation.

Tamppari, L.↗
NASA NTRS1974

An investigation of wave-particle interactions and particle dynamics using electron beams injected from sounding rockets

Electrons with energy up to 40 kV have been injected into semi-trapped orbits from sounding rockets at Wallops Island, Virginia, and at Fort Churchill, Manitoba, Canada. By directing the rocket trajectory to have a horizontal component which in direction and speed matched the bounce displacement of the injected electrons, it was possible to detect conjugate echoes at Wallops and possibly at Churchill, and to study the distribution of the echoes in space, time and energy. By combining observations of many echoes, a composite picture can be obtained of the beam patterns. Atmospheric scattering at the conjugate point of Wallops Island has been extensively studied, and the process produces a scale width of 5-10 m, but with an echo intensity which is only 10% of theoretical estimates.

Winckler, J. R.↗
NASA NTRS1985

Asteroidal source of ordinary chondrites (Meteoritical Society Presidential Address 1984)

The orbital evolution of asteroidal fragments injected into the 3-1 Kirkwood gap resonance at 2.5 AU is investigated on the basis of a Monte Carlo simulation. The diameters of the fragments in the simulation were between 10 cm and 20 km, and it was assumed that the fragments cross the orbital path of the earth every one million years. The effects of close encounter planetary perturbations, the nu dot 6 secular resonance, and the ablative effects of the earth atmosphere were also taken into account. It is found that: (1) the predicted meteorite orbits closely matched the known orbits of ordinary chondrites; and (2) the total flux was in approximate agreement with the observed fall rate of ordinary chondrites. About 90 percent of the predicted impacting bodies were created by fragmentation of larger earth crossing asteroidal fragments, the largest of which were observed in the vicinity of the Apollo-Amor objects. The numerical results are presented in a series of graphs.

Wetherill, G. W.↗
NASA NTRS2000

Flight Results of the Chandra X-ray Observatory Inertial Upper Stage Space Mission

Under contract to NASA, a specially configured version of the Boeing developed Inertial Upper Stage (IUS) booster was provided by Boeing to deliver NASA's 1.5 billion dollar Chandra X-Ray Observatory satellite into a highly elliptical transfer orbit from a Shuttle provided circular park orbit. Subsequently, the final orbit of the Chandra satellite was to be achieved using the Chandra Integral Propulsion System (IPS) through a series of IPS burns. On 23 July 1999 the Shuttle Columbia (STS-93) was launched with the IUS/Chandra stack in the Shuttle payload bay. Unfortunately, the Shuttle Orbiter was unexpectantly inserted into an off-nominal park orbit due to a Shuttle propulsion anomaly occurring during ascent. Following the IUS/Chandra on-orbit deployment from the Shuttle, at seven hours from liftoff, the flight proven IUS GN&C system successfully injected Chandra into the targeted transfer orbit, in spite of the off-nominal park orbit. This paper describes the IUS GN&C system, discusses the specific IUS GN&C mission data load development, analyses and testing for the Chandra mission, and concludes with a summary of flight results for the IUS part of the Chandra mission.

Tillotson, R.↗
NASA NTRS2002

Orbit Determination Support for the Microwave Anisotropy Probe (MAP)

The Microwave Anisotropy Probe (MAP) is a Medium Class Explorers (MIDEX) mission produced in partnership between Goddard Space Flight Center (GSFC) and Princeton University. The main science objective of the MAP mission is to produce an accurate full-sky map of the cosmic microwave background temperature fluctuations anisotropy. MAP was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle at exactly 19:46:46.183 UTC on June 30, 2001. The spacecraft received a nominal direct insertion by the Delta into a 185 km circular orbit. MAP was then maneuvered into a sequence of phasing loops designed to set up a lunar swingby (gravity-assisted acceleration) of the spacecraft onto a transfer trajectory to a Lissajous orbit about the Earth-Sun L2 point. The mission duration is approximately 27 months with 3 to 4 months of transfer time to the final mission orbit about L2. The MAP transfer orbit consisted of 3.5 phasing loops: the first loop has a 7-day period, the second and third loops have a 9-day period, and the last half loop has a 4-day period as illustrated in Figure 1, which also indicates the placement of maneuvers. A Pfinal correction maneuver was performed 18 hours after the last perigee to more closely achieve the targeted lissajous orbit. The lunar encounter or swingby took place approximately 30 days after launch. After the lunar encounter, the spacecraft will cruise for approximately 120 days before it arrives at L2. A Mid-Course Correction (MCC) maneuver was executed seven days after the swingby to further refine the trajectory. Once the MAP satellite is injected into the L2 Lissajous orbit, it will perform occasional stationkeeping maneuvers to maintain the Lissajous orbit for a minimum of two years (and a goal of four years). Because of its complex orbital characteristics, the mission provided a unique challenge to orbit determination (OD) support in many orbital regimes. Extensive trajectory error covariance analysis was performed to predict ephemeris accuracy for the OD process using a Bayesian least-squares technique. The orbit determination error analysis is essential for maneuver planning and maneuver recovery study. Several tracking scenarios were investigated for each phase of the mission. This paper provides a summary of the premission trajectory covariance error analysis, as well as actual real-time OD results. The use and impact of the various tracking stations, systems, and measurements will be discussed. Details of the operational OD support and the inferred OD accuracy will be presented, and the results will be compared to the premission covariance analysis, In addition, there will be a discussion of the challenges presented to OD support including delta-Vs at apogee as well as perigee, and effects due to spacecraft attitude mode, in light of their implications to the OD accuracy and covariance analysis.

Truong, Son H.↗
NASA NTRS1981

Aerobraking mission design - Mission domain and mass performance

The work reported comprises part of an Aerobraking Study that sought to establish a 'feasible mission design, navigation design, and MOS design and to show the desirability of aerobraking for the Venus Orbiting Imaging Radar (VOIR) mission by assessing mission performance, cost, and risk'. The developed software assesses accurately the mass performance of aerobraking and chemical missions that place a spacecraft in orbit about another planet. All injection date/arrival data combinations that provide a trajectory with adequate mass performance are available for further study. The considered analysis has been applied to the Type I trajectories from earth to Venus in 1988. The analysis can easily be applied to missions to Mars and to Titan. The intersection of the determined mass performance domain with the stable orbit domain provides an adequate mission domain for VOIR 1988.

Kerridge, S. J.↗
NASA NTRS2010

Implementation of an Autonomous Multi-Maneuver Targeting Sequence for Lunar Trans-Earth Injection

Using a fully analytic initial guess estimate as a first iterate, a targeting procedure that constructs a flyable burn maneuver sequence to transfer a spacecraft from any closed Moon orbit to a desired Earth entry state is developed and implemented. The algorithm is built to support the need for an anytime abort capability for Orion. Based on project requirements, the Orion spacecraft must be able to autonomously calculate the translational maneuver targets for an entire Lunar mission. Translational maneuver target sequences for the Orion spacecraft include Lunar Orbit Insertion (LOI), Trans-Earth Injection (TEI), and Trajectory Correction Maneuvers (TCMs). This onboard capability is generally assumed to be supplemental to redundant ground computation in nominal mission operations and considered as a viable alternative primarily in loss of communications contingencies. Of these maneuvers, the ability to accurately and consistently establish a flyable 3-burn TEI target sequence is especially critical. The TEI is the sole means by which the crew can successfully return from the Moon to a narrowly banded Earth Entry Interface (EI) state. This is made even more critical by the desire for global access on the lunar surface. Currently, the designed propellant load is based on fully optimized TEI solutions for the worst case geometries associated with the accepted range of epochs and landing sites. This presents two challenges for an autonomous algorithm: in addition to being feasible, the targets must include burn sequences that do not exceed the anticipated propellant load.

Whitley, Ryan J.↗
NASA NTRS2013

Mars Sample Return Using Commercial Capabilities: ERV Trajectory and Capture Requirements

Mars Sample Return was presented as the highest priority planetary science mission of the next decade [1]. Lemke et al. [2] present a Mars Sample Return mission concept in which the sample is returned directly from the surface of Mars to an Earth orbit. The sample is recovered in Earth Orbit instead of being transferred between spacecraft in Mars Orbit. This paper provides the details of this sample recovery in Earth orbit and presents as such a sub-element of the overall Mars sample return concept given in [2]. We start from the assumption that a Mars Ascent Vehicle (MAV), initially landed on Mars using a modified SpaceX Dragon capsule, has successfully delivered the sample, already contained within an Earth Return Vehicle (ERV), to a parking orbit around Mars. From the parking orbit, the ERV imparts sufficient Delta-V to inject itself into an earthbound trajectory and to be captured into an Earth orbit eventually. We take into account launch window and Delta-V considerations as well as the additional constraint of increased safety margins imposed by planetary protection regulations. We focus on how to overcome two distinct challenges of the sample return that are driven by the issues of planetary protection: (1) the design of an ERV trajectory meeting all the requirements including the need to avoid contamination of Earth's atmosphere; (2) the concept of operations for retrieving the Martian samples in Earth orbit in a safe way. We present an approach to retrieve the samples through a rendezvous between the ERV and a second SpaceX Dragon capsule. The ERV executes a trajectory that brings it from low Mars orbit (LMO) to a Moon-trailing Earth orbit at high inclination with respect to the Earth-Moon plane. After a first burn at Trans-Earth Injection (TEI), the trajectory uses a second burn at perigee during an Earth flyby maneuver to capture the ERV in Earth orbit. The ERV then uses a non-propulsive Moon flyby to come to a near-circular Moon-trailing orbit. To perform the Earth Orbit Rendezvous (EOR), a second Dragon capsule is then launched from Earth and a similar lunar flyby is performed to rendezvous with the ERV. The requirements for rendezvous, close proximity operations and capture of the sample canister are described. A concept of operations for sample retrieval is presented along with design specifications of the ERV, the required modifications to the Dragon capsule, as well as the hardware, software, sensors, actuators, and capture mechanisms used. In our concept, a container is mounted to the front hatch of Dragon, capable of accommodating the sample canister and sealing it from the rest of the capsule. The sample canister is captured using a robotic arm with a magnetic grappling mechanism. Dragon then performs a propulsive maneuver to return to Earth for a controlled re-entry while the ERV (sans sample container) is left in the Moon trailing orbit. Contingency cases and related mitigation strategies are also discussed, including the advantages and disadvantages of performing the ERV rendezvous with a crew.

Faber, Nicolas F.↗
NASA NTRS1973

Pioneer Venus mission plan for atmospheric probes and an orbiter.

The mission concept is described for a spin-stabilized spacecraft being designed to deliver an array of aerodynamic probes over the surface of Venus, and, on a second flight mission, to inject the spacecraft into a Venusian orbit. Large and Small Probes and the instrumented 'Bus' structure will enter the Venusian atmosphere and directly measure its characteristics. An orbiting configuration of the same basic Bus spacecraft will serve as a platform from which to examine the planetary environment, the outer atmosphere and its interaction with the solar wind, as well as surface topography. The orbital mission will extend over a complete revolution of the planet around the sun and around its own axis. Requirements for cost effectiveness and firm adherence within early cost projections are emphatic program guidelines which affect every major element of the mission plan and conceptual design.

Dyer, J. W.↗
NASA NTRS2000

Artificial Gravity for Mars Missions: The Different Design and Development Options

One of the major impediments to human Mars missions is the development of appropriate countermeasures for long term physiological response to the micro-gravity environment. A plethora of countermeasure approaches have been advanced from strictly pharmacological measures to large diameter rotating spacecraft that would simulate a 1-g environment (the latter being the most conservative from a human health perspective). The different approaches have significantly different implications not only on the overall system design of a Mars Mission Vehicle (MMV) but on the necessary earth-orbiting platform that would be required to qualify the particular countermeasure system. it is found that these different design options can be conveniently categorized in terms of the order of magnitude of the rotation diameter required (100's, 10's, 1's, 0 meters). From this, the different mass penalties associated with each category can be generally compared. The overall objective of the countermeasure system should be to maximize crew safety and comfort, minimize exercise protocol time (i.e., the time per day that each crew member would have to participate in the exercise/countermeasure), maximize countermeasure effectiveness, and minimize the associated system mass penalty of the Mars Mission Vehicle (in terms of fraction of IMLEO - Injected Mass in Low Earth Orbit).

Murbach, Marcus↗
NASA NTRS1991

The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

The following paper reports on a design study of a novel space transportation concept known as a 'NIMF' (Nuclear rocket using Indigenous Martian Fuel). The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO2. This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34,000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface.

Zubrin, Robert M.↗
DOE OSTI2025

Coupling of Nondegenerate Topological Modes in Nitrogen Core-Doped Graphene Nanoribbons

Nitrogen core-doping of graphene nanoribbons (GNRs) allows trigonal planar carbon atoms along the backbone of GNRs to be substituted by higher-valency nitrogen atoms. The excess valence electrons are injected into the π-orbital system of the GNR, thereby changing not only its electronic occupation but also its topological properties. We have observed this topological change by synthesizing dilute nitrogen core-doped armchair GNRs with a width of five atoms (N 2 -5-AGNRs). The incorporation of pairs of trigonal planar nitrogen atoms results in the emergence of topological boundary states at the interface between doped and undoped segments of the GNR. These topological boundary states are offset in energy by approximately ΔE = 300 meV relative to the topological end states at the termini of finite 5-AGNRs. Scanning tunneling microscopy (STM) and spectroscopy (STS) reveal that for finite GNRs the two types of topological states can interact through a linear combination of orbitals, resulting in a pair of asymmetric hybridized states. This behavior is captured by an effective Hamiltonian of nondegenerate diatomic molecules, where the analogous interatomic hybridization interaction strength is tuned by the distance between GNR topological modes.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1962

Empire. a study of early manned interplanetary missions final report, may 26 - nov. 25, 1962

This report summarizes the investigations and results of the EMPIRE Study Program undertaken by Aeronutronic Division of Ford Motor Company for the Future Projects Office, Marshall Space Flight Center, under Contract NAS8-5025. The dual planet flyby missions of the Crocco and Symmetric trajectory classes are discussed. The Crocco mission with an August 1971 launch window requires an interplanetary injection velocity increment of i0.i km/sec, has a return velocity of 13.5 km/sec, and takes approximately 400 days. The Symmetric mission with a July 1970 launch window has an injection velocity increment of 5.3 km/sec, a return velocity of 15.8 km/sec, and takes approximately 630 days. Additional results of the trajectory studies and abort trajectories are reported. The guidance and navigation subsystem, midcourse corrections, and planetary approach corrections are discussed. A detailed analysis of the reentry phase of EMPIRE includes consideration of an Apollo-type, a Drag Brake, and a lifting-type reentry vehicle to return the six-man crew at mission completion or in an aborted condition. The High L/D reentry vehicle is used in the missions considered. The various technological areas required for design criteria are developed and several spacecraft designs are considered. The all chemical propulsion Crocco system is discarded due to weight, complexity, and cost. The nuclear injected Crocco is treated in a similar manner. The lower energy injection for the Symmetric Mission leads to the feasibility of a nuclear injected vehicle with an Earth orbit weight of about 180,000 kilograms (400,000 pounds) before interplanetary transit. In addition, two chemical symmetric vehicles are treated. Conservative radiation exposures are derived, for the 630 day mission, of less than 200 REM and a polyethelene radiation shelter is designed. Scientific aspects of the missions are discussed. Mission Success Probabilities are presented for the various missions considered and for Saturn C-5, Nova, and Super-Nova Earth launch vehicles in light of possible development. The need for acceleration of nuclear rocket engine developments and auxiliary power developments is indicated. Definition of a larger nuclear engine of the order of 200,000 pounds thrust and about 800 seconds burning time or 50,000 pound thrust and 3600 seconds burning time is indicated for the Symmetric Mission in 1970 (energy requirements are higher in 1972 and for later launch due to the less favorable position of Mars)_ Immediate development of this advanced nuclear propulsion capability is recommended. A Development Plan and Funding Schedule is given for the 1970 launch window pinpointing the critical development areas and indicating a total program cost of $12.6 billion independent of other programmed R&D costs. In conclusion, technological feasibility for an early manned dual planet Mars-Venus flyby is believed to be demonstrated in this study. Several areas of accelerated development and experimental confirmation of theory are pinpointed. The necessary funding and development of Nova or orbital operations capability with Saturn C-5's is required. The 1970 launch window appears to offer the least expensive Symmetric Mission for several years into the 1980's.

F. P. Dixon↗
NASA NTRS1972

Atlas/Centaur-29 Intelsat IV (F5) flash flight report

ATLAS/CENTAUR-29 was launched from ETR Complex 36B, June 13, 1972, at a flight azimuth of 101 degrees. The ETR test number was 1240. The launch vehicle consisted of an ATLAS SLV-3C first stage and a CENTAUR second stage. The spacecraft was INTELSAT 4 (F-5). The launch occurred at the opening of the window. Indications are that the spacecraft was injected into the required transfer orbit and has achieved it's final orbit. All orbital parameters were extremely close to nominal. A nominal apogee motor burn took place at 2101 EDT on June 14, 1972 at 124 degrees East longitude and the final duty station of the spacecraft will be at 62 degrees East longitude. A two week communications checkout was started on or about June 15, 1972.

Source record↗
NASA NTRS1980

A solar activity monitoring platform for SCADM

The adaptation of proven space probe technology is proposed as a means of providing a solar activity monitoring platform which could be injected behind the Earth's orbital position to give 3 to 6 days advanced coverage of the solar phenomenon on the backside hemisphere before it rotates into view and affects terrestrial activities. The probe would provide some three dimensional discrimination within the ecliptic latitude. This relatively simple off-Earth probe could provide very high quality data to support the SCADM program, by transmitting both high resolution video data of the solar surface and such measurements of solar activity as particle, X-ray, ultraviolet, and radio emission fluxes. Topics covered include the orbit; constraints on the spacecraft; subsystems and their embodiments; optical imaging sensors and their operation; and the radiation-pressure attitude control system are described. The platform would be capable of mapping active regions on an hourly basis with one arc-second resolution.

Kissell, K. E.↗
NASA NTRS1982

Nuclear electric propulsion mission to Neptune

A nuclear electric propulsion (NEP) mission to Neptune is studied. Results of trajectory and propulsion parameter optimization are presented which lead to a selection of key electric propulsion system parameters. Spiral earth escape and spiral capture modes are assumed in the optimization studies. Flight time vs payload for Space Transportation System (STS)/Centaur injection and Triton-aided orbit capture at Neptune are also briefly investigated as methods of improving mission performance. Radiation effects of the earth spiral phase are discussed. The effects of NEP technology on science payloads and mission and system designs are evaluated. NEP is shown to be a very flexible and high performing propulsion technology.

Nock, K. T.↗
NASA NTRS1992

Space Exploration Initiative - Overview, current status, near term plans

An overview is presented of the Space Exploration Initiative (SEI) describing current activities, accomplishemnts, and directions. The program is considered in terms of activities relevant to the three phases which are: manned robotic/precursor missions, manned lunar missions, and manned Mars missions. SEI is shown to be in the First Lunar Outpost (FLO) mission phase which is based on an early manned lunar return, and development is described of earth-to-orbit (ETO) vehicles, a translunar injection stage, and a lunar surface habitat. Two ETO vehicles are considered, and the translunar injection stage is illustrated. The FLO mission is also considered as an initiation of the design activity for the Mars mission so that the requirements and preliminary designs could be defined. The FLO program demonstrates the feasibility of a manned return to the moon as an initiation of subsequent SEI activities.

Adams, Alan M.↗
NASA NTRS2001

Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity, but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to-diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L/D vortex chamber concept for gel propellants was conducted by Michaels. He used both triplet (two oxidizer and one fuel orifices) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called the subject injection scheme as Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 92%, can be obtained. MSFC and the U.S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen/hydrocarbon propellant system. This vortex chamber concept is currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX)/hydrocarbon fuel (RPM) system has been derived from the one for the gel propellant.

Trinh, Huu P.↗