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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Nuclear concepts/propulsion

Nuclear thermal and nuclear electric propulsion systems will enable and/or enhance important space exploration missions to the moon and Mars. Current efforts are addressing certain research areas, although NASA and DOE still have much work yet to do. Relative to chemical systems, nuclear thermal propulsion offers the potential of reduced vehicle weight, wider launch windows. and shorter transit times, even without aerobrakes. This would improve crew safety by reducing their exposure to cosmic radiation. Advanced materials and structures will be an important resource in responding to the challenges posed by safety and test facility requirements, environmental concerns, high temperature fuels and the high radiation, hot hydrogen environment within nuclear thermal propulsion systems. Nuclear electric propulsion (NEP) has its own distinct set of advantages relative to chemical systems. These include low resupply mass, the availability of large amounts of onboard electric power for other uses besides propulsion, improved launch windows, and the ability to share technology with surface power systems. Development efforts for NEP reactors will emphasize long life operation of compact designs. This will require designs that provide high fuel burnup and high temperature operation along with personnel and environmental safety.

Miller, Thomas J.↗
NASA NTRS1990

An unmanned probe to Pluto

Now that Voyager II has completed its grand tour of the solar system, all the planets in the solar system, with the exception of Pluto, have been studied. Even now, missions to return to Mercury, Venus, Mars Jupiter, and Saturn are currently flying or are planned. However, a mission to explore Pluto is not, at the present time, being considered seriously. The design problem presented to the students was very general, i.e., design an unmanned mission to Pluto with a launch window constraint of the years 2000 to 2010. All other characteristics of the mission, such as mission type (flyby, orbiter, lander, penetrator), scientific objectives and payload, and the propulsion system were to be determined by the design teams. The design studies exposed several general problems to be solved. Due to the extreme distance to Pluto (and a corresponding travel time in the range of 10 to 25 years), the spacecraft had to be lighter and more robust than current spacecraft designs. In addition, advanced propulsion concepts had to be considered. These included the new generation of launch vehicles and upper stages and nuclear electric propulsion. The probe design offered an abundance of synthesis and analysis problems. These included sizing trade studies, selection of subsystem components, analysis of spacecraft dynamics, stability and control, structural design and material selection, trajectory design, and selection of scientific equipment. Since the characteristics of the mission, excluding the launch window, were to be determined by the design teams, the solutions varied widely.

Source record↗
NASA NTRS1975

Non-numeric computation for high eccentricity orbits

Geocentric orbits of large eccentricity (e = 0.9 to 0.95) are significantly perturbed in cislunar space by the sun and moon. The time-history of the height of perigee, subsequent to launch, is particularly critical. The determination of 'launch windows' is mostly concerned with preventing the height of perigee from falling below its low initial value before the mission lifetime has elapsed. Between the extremes of high accuracy digital integration of the equations of motion and of using an approximate, but very fast, stability criteria method, this paper is concerned with the developement of a method of intermediate complexity using non-numeric computation. The computer is used as the theory generator to generalize Lidov's theory using six osculating elements. Symbolic integration is completely automatized and the output is a set of condensed formulae well suited for repeated applications in launch window analysis. Examples of applications are given.

Sridharan, R.↗
NASA NTRS1999

Performance Optimization of the Gasdynamic Mirror Propulsion System

Nuclear fusion appears to be a most promising concept for producing extremely high specific impulse rocket engines. Engines such as these would effectively open up the solar system to human exploration and would virtually eliminate launch window restrictions. A preliminary vehicle sizing and mission study was performed based on the conceptual design of a Gasdynamic Mirror (GDM) fusion propulsion system. This study indicated that the potential specific impulse for this engine is approximately 142,000 sec. with about 22,100 N of thrust using a deuterium-tritium fuel cycle. The engine weight inclusive of the power conversion system was optimized around an allowable engine mass of 1500 Mg assuming advanced superconducting magnets and a Field Reversed Configuration (FRC) end plug at the mirrors. The vehicle habitat, lander, and structural weights are based on a NASA Mars mission study which assumes the use of nuclear thermal propulsion' Several manned missions to various planets were analyzed to determine fuel requirements and launch windows. For all fusion propulsion cases studied, the fuel weight remained a minor component of the total system weight regardless of when the missions commenced. In other words, the use of fusion propulsion virtually eliminates all mission window constraints and effectively allows unlimited manned exploration of the entire solar system. It also mitigates the need to have a large space infrastructure which would be required to support the transfer of massive amounts of fuel and supplies to lower a performing spacecraft.

Emrich, William J., Jr.↗
NASA NTRS2011

Strategic Implications of Human Exploration of Near-Earth Asteroids

The current United States Space Policy [1] as articulated by the White House and later confirmed by the Congress [2] calls for [t]he extension of the human presence from low-Earth orbit to other regions of space beyond low-Earth orbit will enable missions to the surface of the Moon and missions to deep space destinations such as near-Earth asteroids and Mars. Human exploration of the Moon and Mars has been the focus of numerous exhaustive studies and planning, but missions to Near-Earth Asteroids (NEAs) has, by comparison, garnered relatively little attention in terms of mission and systems planning. This paper examines the strategic implications of human exploration of NEAs and how they can fit into the overall exploration strategy. This paper specifically addresses how accessible NEAs are in terms of mission duration, technologies required, and overall architecture construct. Example mission architectures utilizing different propulsion technologies such as chemical, nuclear thermal, and solar electric propulsion were formulated to determine resulting figures of merit including number of NEAs accessible, time of flight, mission mass, number of departure windows, and length of the launch windows. These data, in conjunction with what we currently know about these potential exploration targets (or need to know in the future), provide key insights necessary for future mission and strategic planning.

Drake, Bret G.↗
NASA NTRS2023

Managing the Lucy Mission

Lucy is NASA’s 13th Discovery-class mission. It will be the first mission to encounter seven Trojan asteroids in a series of five flybys. Lucy is led by Dr. Hal Levison of the Southwest Research Institute. As the Principal Investigator, Dr. Levison is accountable to NASA for the success of the investigation, with full responsibility for its scientific integrity and execution within committed cost and schedule. Day to day management of the Lucy project is delegated to Goddard Space Flight Center. During Lucy’s development, the project experienced a once-in-a-lifetime pandemic that threatened the team’s ability to meet the planetary launch window. Despite this challenge, Lucy launched on time and under budget. This chapter discusses key factors that enabled project success.

Project Management↗
NASA NTRS1999

Launch Collision Probability

This report analyzes the probability of a launch vehicle colliding with one of the nearly 10,000 tracked objects orbiting the Earth, given that an object on a near-collision course with the launch vehicle has been identified. Knowledge of the probability of collision throughout the launch window can be used to avoid launching at times when the probability of collision is unacceptably high. The analysis in this report assumes that the positions of the orbiting objects and the launch vehicle can be predicted as a function of time and therefore that any tracked object which comes close to the launch vehicle can be identified. The analysis further assumes that the position uncertainty of the launch vehicle and the approaching space object can be described with position covariance matrices. With these and some additional simplifying assumptions, a closed-form solution is developed using two approaches. The solution shows that the probability of collision is a function of position uncertainties, the size of the two potentially colliding objects, and the nominal separation distance at the point of closest approach. ne impact of the simplifying assumptions on the accuracy of the final result is assessed and the application of the results to the Cassini mission, launched in October 1997, is described. Other factors that affect the probability of collision are also discussed. Finally, the report offers alternative approaches that can be used to evaluate the probability of collision.

Bollenbacher, Gary↗
NASA NTRS2019

Ion Propulsion for Solar System Exploration

NASA's Evolutionary Xenon Thruster (NEXT) is ready for transition-to-flight. The thruster has completed all qualification-level environmental testing, and has demonstrated a xenon propellant throughput, total impulse, and total operating hours greatly in excess of anticipated planetary science mission requirements, and exceeding that achieved by any other thruster technology in the history of electric propulsion. NEXT is the next generation system, a natural progression in technology from that implemented successfully on the Deep-Space one and Dawn missions, developed at NASA's Glenn Research Center in Cleveland, Ohio. The first implementation of NEXT will be on NASA 's Double Asteroid Redirection Test (DART). DART will be the first demonstration of the kinetic impact technique to change the motion of an asteroid in space. The DART mission is in Phase C, led by Johns Hopkins University Applied Physics Laboratory. The DART spacecraft will utilize the NASA Evolutionary Xenon Thruster solar electric propulsion system as its primary in-space propulsion system. By utilizing NEXT, DART is able to gain significant flexibility to the mission timeline and launch window, as well as decrease in launch vehicle cost. This presentation will review NASA's investment strategy in electric propulsion _ in particular gridded ion thruster technology _ as it applies to solar system exploration. Results obtained from implementing this technology on Deep-Space one and Dawn will be reviewed. Mission studies which highlight the impacts of the NEXT technology will be discussed, and near-term proposed and scheduled missions including DART and CAESAR (Comet Astrobiology Exploration Sample Return) will be reviewed.

Electric Propulsion↗
NASA NTRSDate not supplied

Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey recently identified Uranus and Neptune - called Ice Giants - as the priority destinations for science[1]. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission will deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a long cruise time to destination (between 12 and 15 years) and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture is a method of orbital control that uses aerodynamic forces generated on a vehicle by the planet’s atmosphere to modulate a spacecraft’s trajectory, allowing mission designers to target the final orbital state. For the Uranus mission, using aerocapture for orbital insertion can decrease not only the cruise time to the destination by 2 - 3 years, but the propellant required to achieve orbital insertion (by more than 40%) which would, in turn, increase the available science payload and reduce the timeline for retrieving data vital to the mission’s science objectives[2]. Achieving orbital insertion via aerocapture requires novel algorithms for Guidance, Navigation and Control[3], and mass-efficient Thermal Protection Systems (TPS) performing in a new atmosphere. This paper will focus on the selection and tailoring of the Thermal Protection Systems for the forebody and aftbody heatshields of an aerocapture mission to Uranus. While preliminary results indicate that multiple systems in NASA’s repertoire are capable of performing in the predicted aerothermal environment there are unique aspects like the inert environment that affect ablation efficiency, and the heatload for aerocapture trajectories to the outer planets are among the highest of any mission to-date[4]. These two factors may impose operational requirements to heatshield separation in order to minimize thermal soak to the payload, and may demand TPS thickness and configurations that have not yet been demonstrated. This paper will discuss the updated maturity, manufacturing, and performance capabilities of candidate thermal protection systems, with specific areas of need highlighted to make thermal protection systems viable for use in the recommended Uranus Orbiter and Probe mission.

Uranus↗
NASA NTRSDate not supplied

Improvements to Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey identified Uranus and Neptune - called Ice Giants - as the priority destinations for science. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission concept planned to deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a cruise phase of at least 12 years and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture uses aerodynamic forces generated on a vehicle by the planet's atmosphere to modulate a spacecraft's trajectory, decreasing spacecraft velocity, and allowing mission designers to target the final orbital state. Aerocapture reduces the time-of-flight from Earth to Uranus over a fully propulsive solution, opening up more launch opportunities to arrive in the 2040's to the mission's science objectives. Aerocapture also allows a payload mass increase by mitigating the need for fuel to retropropulsively insert the payload into orbit, thereby increase the science that can be performed. For an aerocapture mission structure using a traditional aeroshell to deliver the UOP scientific payload to Uranus, Conformal Phenolic Impregnated Carbon Ablator (C-PICA) was determined to be the best-performing forebody thermal protection system (TPS) and other candidate aftbody TPS options were presented as feasible. This paper focuses on A) evaluating C-PICA as a forebody TPS using stressing entry conditions associated with a large range of potential Uranus flagship launch vehicles and interplanetary trajectories, B) widening the aftbody TPS candidates for new mass-efficient and cost-efficient solutions, and C) a summary list of actions remaining to provide a technically feasible and supply-robust set of TPS for an aerocapture vehicle to the Ice Giants is presented.

C-PICA↗
NASA NTRSDate not supplied

Improvements to Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey identified Uranus and Neptune - called Ice Giants - as the priority destinations for science. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission concept planned to deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a cruise phase of at least 12 years and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture uses aerodynamic forces generated on a vehicle by the planet's atmosphere to modulate a spacecraft's trajectory, decreasing spacecraft velocity, and allowing mission designers to target the final orbital state. Aerocapture reduces the time-of-flight from Earth to Uranus over a fully propulsive solution, opening up more launch opportunities to arrive in the 2040's to the mission's science objectives. Aerocapture also allows a payload mass increase by mitigating the need for fuel to retropropulsively insert the payload into orbit, thereby increase the science that can be performed. For an aerocapture mission structure using a traditional aeroshell to deliver the UOP scientific payload to Uranus, Conformal Phenolic Impregnated Carbon Ablator (C-PICA) was determined to be the best-performing forebody thermal protection system (TPS) and other candidate aftbody TPS options were presented as feasible. This paper focuses on A) evaluating C-PICA as a forebody TPS using stressing entry conditions associated with a large range of potential Uranus flagship launch vehicles and interplanetary trajectories, B) widening the aftbody TPS candidates for new mass-efficient and cost-efficient solutions, and C) a summary list of actions remaining to provide a technically feasible and supply-robust set of TPS for an aerocapture vehicle to the Ice Giants is presented.

C-PICA↗
NASA NTRS1973

Summary of Saturn swingby missions to Uranus

The interplanetary trajectory characteristics for missions to Uranus, which employ an intermediate swingby of Saturn to reduce the total trip time are summarized. Opportunities for such swingby missions will occur from 1979 through 1987 and not again until 2025. The general trajectory characteristics (C sub 3; departure, swingby, and arrival dates; swingby radius; and arrival speed) are evaluated, and payload and launch window information for a Titan 3E/Centaur/TE-364-4 class launch vehicle is provided.

Manning, L. A.↗
NASA NTRS1973

Mission Planning for Pioneer Saturn/Uranus Atmospheric Probe Missions

Mission planning for a series of atmospheric probe missions to Saturn and Uranus using a modified Pioneer spacecraft launched in 1979 and 1980 was examined. The operational options and the associated systems requirements consistent with the major scientific goals and spacecraft constraints of the missions is summarized. It is feasible to obtain in-situ atmospheric measurements in the atmosphere of Saturn and Uranus down to a pressure level of 10 bars using a common probe and spacecraft design. Spacecraft can be launched to both objectives with an adequate launch window in 1979 and 1980 using a Titan/Centaur launch vehicle with a TE-364-4 upper stage. Other scientific objectives can be accomplished by the flyby spacecraft. Encounters with the satellite Titan and RF occultations of Saturn, the ring system of Saturn, and Uranus can be obtained.

Swenson, B. L.↗
NASA NTRS1983

Performance requirements analysis for payload delivery from a space station

Operations conducted from a space station in low Earth orbit which have different constraints and opportunities than those conducted from direct Earth launch were examined. While a space station relieves many size and performance constraints on the space shuttle, the space station's inertial orbit has different launch window constraints from those associated with customary Earth launches which reflect upon upper stage capability. A performance requirements analysis was developed to provide a reference source of parametric data, and specific case solutions and upper stage sizing trade to assist potential space station users and space station and upper stage developers assess the impacts of a space station on missions of interest.

Friedlander, A. L.↗
NASA NTRS1990

Space vehicle deployment from Space Station

When launching a spacecraft from Earth parking orbit to deep space, it is highly desirable to have the hyperbolic excess velocity vector (v-infinity) contained in the parking orbit plane. Ground launches can force the parking orbit plane to contain the v-infinity vector by using launch azimuth and lift-off time as independent variables. When launching from the Space Station, a new set of variables comes into play. The Space Station orbit is of fixed inclination but precessing due to the Earth's oblateness. Its plane will seldom (and may never) contain the desired v-infinity vector. Consequently, the departure strategy will usually require multiple burns and include a plane change. Also, the concept of 'launch window' will be somewhat different from Earth surface launches. An analysis of the deployment of interplanetary spacecraft from the Space Station is described, with emphasis on the effect of the trajectory characteristics on station operations. Several planetary mission types are analyzed, including manned Mars missions and unmanned high declination departures. The constraint of Space Station orbit nodal position is quantified and the operational implications for Space Station reboost strategy are examined.

Henry, Paul K.↗
NASA NTRS1997

Sounding of the Cleft Ion Fountain Energization Region

The objectives of the ground-based observations in support of the SCIFER are: Acquire and display ionospheric conditions prior to launch to aid in the establishment of launch criteria in real time. Observers at both stations participated in real-time visual interpretation. Solar wind data from IMP-8 and WIND were acquired and interpreted in real time. Telephonic and data links were established at the observatory for the launch window period. Ground-based observatory countdown and launch criteria were developed. 2) Relate optical and magnetic ionospheric signatures observed from the ground to magnetospheric boundaries in the energetic particle flux measured at the payload. The energetic electron trapping boundary was found to correspond to the equatorward edge of the discrete auroral arcs forming the dayside aurora. The energetic electron trapping boundary was found to correspond to the poleward edge of pulsating aurora. The pulsating aurora was found to correspond to one second bursts of energy-dispersed electrons originating in the equatorial plane. Pulsations at larger intervals corresponded to travel times to the conjugate region and return. The pulsating aurora was also directly linked to the geomagnetic pulsations and traveling magnetic vortices, all occurring equatorward of the trapping boundary. 630 nm emission corresponding to less than 10 eV electron precipitation was observed equatorward of the trapping boundary (L=15) and ascribed to photoelectrons from the sunlit conjugate region. 3) Aid in the interpretation of time/space incongruities in the rocket data. The motion of the payload conjugate across the aurora showed that the payload passed over three distinct arc systems on the poleward side of the trapping boundary. These results were reported in a series of articles to be printed in Geophysical Research Letters on June 15, l996.

Kintner, P.↗