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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 415 records · Page 23

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NASA NTRS1992

How shall we go to Mars? A review of mission scenarios

Four scenarios for initial Mars missions are examined in this overview of strategies. The scenarios are compared to determine the most effective option with respect to the effects of reduced gravity, exposure to space radiation, and cost. Mission cost is evaluated based on initial mass in earth orbit for the four mission types: classes-opposition, conjunction, fast-transfer conjunction, and split sprint. The reduced gravitational force on the Martian surface is considered to be a critical factor because the physical deconditioning at 3/8 earth gravity is assumed to approximate that of zero gravity. Exposure to radiation is not determined to be a definite discriminator because none of the scenarios exceed established limits. The split sprint is championed based on the high energy transfers considered and with chemical propulsion (or preferably nuclear thermal propulsion) and aerobraking.

Walberg, Gerald↗

NASA NTRS1992

Engine system assessment study using Martian propellants

A feasibility study was performed that identified and characterized promising chemical propulsion system designs that utilize two or more of the propellant combinations: LOX/H2, LOX/CH4 and LOX/CO. The engine systems examined focused on the usage of common subsystem/component hardware where feasible. From the evaluation baseline employed, tripropellant MTV LOX cooled and bipropellant LEV and MEV engine systems are identified.

Pelaccio, D.↗

NASA NTRS1992

Engine system assessment study using Martian propellants

A top-level feasibility study was conducted that identified and characterized promising chemical propulsion system designs which use two or more of the following propellant combinations: LOX/H2, LOX/CH4, and LOX/CO. The engine systems examined emphasized the usage of common subsystem/component hardware where possible. In support of this study, numerous mission scenarios were characterized that used various combinations of Earth, lunar, and Mars propellants to establish engine system requirements to assess the promising engine system design concept examined, and to determine overall exploration leverage of such systems compared to state-of-the-art cryogenic (LOX/H2) propulsion systems. Initially in the study, critical propulsion system technologies were assessed. Candidate expander and gas generator cycle LOX/H2/CO, LOX/H2/CH4, and LOX/CO/CH4 engine system designs were parametrically evaluated. From this evaluation baseline, tripropellant Mars Transfer Vehicle (MTV) LOX cooled and bipropellant Lunar Excursion Vehicle (LEV) and Mars Excursion Vehicle (MEV) engine systems were identified. Representative tankage designs for a MTV were also investigated. Re-evaluation of the missions using the baseline engine design showed that in general the slightly lower performance, smaller, lower weight gas generator cycle-based engines required less overall mission Mars and in situ propellant production (ISPP) infrastructure support compared to the larger, heavier, higher performing expander cycle engine systems.

Pelaccio, Dennis↗

NASA NTRS1995

Aerospace Environmental Technology Conference: Exectutive summary

The mandated elimination of CFC's, Halons, TCA, and other ozone depleting chemicals and specific hazardous materials has required changes and new developments in aerospace materials and processes. The aerospace industry has been involved for several years in providing product substitutions, redesigning entire production processes, and developing new materials that minimize or eliminate damage to the environment. These activities emphasize replacement cleaning solvents and their application verifications, compliant coatings including corrosion protection systems, and removal techniques, chemical propulsion effects on the environment, and the initiation of modifications to relevant processing and manufacturing specifications and standards. The papers from this conference are being published in a separate volume as NASA CP-3298.

Whitaker, A. F.↗

NASA NTRS1995

Aerospace Environmental Technology Conference

The mandated elimination of CFC's, Halons, TCA, and other ozone depleting chemicals and specific hazardous materials has required changes and new developments in aerospace materials and processes. The aerospace industry has been involved for several years in providing product substitutions, redesigning entire production processes, and developing new materials that minimize or eliminate damage to the environment. These activities emphasize replacement cleaning solvents and their application verifications, compliant coatings including corrosion protection systems, and removal techniques, chemical propulsion effects on the environment, and the initiation of modifications to relevant processing and manufacturing specifications and standards. The Executive Summary of this Conference is published as NASA CP-3297.

Whitaker, A. F.↗

NASA NTRS1995

University of Tennessee Center for Space Transportation and Applied Research (CSTAR)

The Center for Space Transportation and Applied Research had projects with space applications in six major areas: laser materials processing, artificial intelligence/expert systems, space transportation, computational methods, chemical propulsion, and electric propulsion. The closeout status of all these projects is addressed.

Source record↗

NASA NTRS1993

Center for Space Transportation and Applied Research Fifth Annual Technical Symposium Proceedings

This Fifth Annual Technical Symposium, sponsored by the UT-Calspan Center for Space Transportation and Applied Research (CSTAR), is organized to provide an overview of the technical accomplishments of the Center's five Research and Technology focus areas during the past year. These areas include chemical propulsion, electric propulsion, commerical space transportation, computational methods, and laser materials processing. Papers in the area of artificial intelligence/expert systems are also presented.

Source record↗

NASA NTRS1996

Chemical Microthruster Options

Chemical propulsion systems with potential application to microsatellites are classified by propellant phase, i.e. gas, liquid, or solid. Four promising concepts are selected based on performance, weight, size, cost, and reliability. The selected concepts, in varying stages of development, are advanced monopropellants, tridyne(TM), electrolysis, and solid gas generator propulsion. Tridyne(TM) and electrolysis propulsion are compared vs. existing cold gas and monopropellant systems for selected microsatellite missions. Electrolysis is shown to provide a significant weight advantage over monopropellant propulsion for an orbit transfer and plane change mission. Tridyne(TM) is shown to provide a significant advantage over cold gas thrusters for orbit trimming and spacecraft separation.

DeGroot, Wim↗

NASA NTRS1997

Second Aerospace Environmental Technology Conference

Whitaker, A. F.↗

NASA NTRS1997

Second Aerospace Environmental Technology Conference

The mandated elimination of CFC'S, Halons, TCA, and other ozone depleting chemicals and specific hazardous materials has required changes and new developments in aerospace materials and processes. The aerospace industry has been involved for several years in providing product substitutions, redesigning entire production processes, and developing new materials that minimize or eliminate damage to the environment. These activities emphasize replacement cleaning solvents and their application, verification, compliant coatings including corrosion protection system and removal techniques, chemical propulsion effects on the environment, and the initiation of modifications to relevant processing and manufacturing specifications and standards.

A F Whitaker↗

NASA NTRS1999

Miniature Heat Transport System for Nanosatellite Technology

The scientific understanding of key physical processes between the Sun and the Earth require simultaneous measurements from many vantage points in space. Nano-satellite technologies will enable a class of constellation missions for the NASA Space Science Sun-Earth Connections. This recent emphasis on the implementation of smaller satellites leads to a requirement for development of smaller subsystems in several areas. Key technologies under development include: advanced miniaturized chemical propulsion; miniaturized sensors; highly integrated, compact electronics; autonomous onboard and ground operations; miniatures low power tracking techniques for orbit determination; onboard RF communications capable of transmitting data to the ground from far distances; lightweight efficient solar array panels; lightweight, high output battery cells; lightweight yet strong composite materials for the nano-spacecraft and deployer-ship structures. These newer smaller systems may have higher power densities and higher thermal transport requirements than seen on previous small satellites. Furthermore, the small satellites may also have a requirement to maintain thermal control through extended earth shadows, possibly up to 8 hours long. Older thermal control technology, such as heaters, thermostats, and heat pipes, may not be sufficient to meet the requirements of these new systems. Conversely, a miniature two-phase heat transport system (Mini-HTS) such as a Capillary Pumped Loop (CPL) or Loop Heat Pipe (LBP) is a viable alternative. A Mini-HTS can provide fine temperature control, thermal diode action, and a highly efficient means of heat transfer. The Mini-HTS would have power capabilities in the range of tens of watts or less and provide thermal control over typical spacecraft ranges. The Mini-HTS would allow the internal portion of the spacecraft to be thermally isolated from the external radiator, thus protecting the internal components from extreme cold temperatures during an eclipse. The Mini-HTS would transport the beat from these components to a radiator during their operational modes, and it would be shutdown during non-operational or eclipse modes. Shutdown of the Mini-HTS would be accomplished with small heaters and has been successfully demonstrated on numerous occasions, both in the lab and on flight experiments. Efforts are now underway to miniaturize two-phase heat transport systems for the Nanosatellite project, with potential application to other small satellite programs. 'ne goal of this project is to design, build, and test miniature heat transport systems (MHTS) that would demonstrate the feasibility of a small Capillary Pumped Loop (CPL) or Loop Heat Pipe (LBP).

Douglas, Donya M,↗

NASA NTRS2000

Solar Sail Trajectories for Solar Polar and Heliopause Missions

Over the last several years, interest in a more vigorous space exploration program has renewed interest in the use of solar sails for the more demanding space missions. Solar sail is eminently suited for some of the higher energy missions since no fuel is consumed and the only criteria is that of the total time required to perform a particular mission. Although solar sail missions to planets and small bodies have been examined previously, and reported in the literature' classes of space missions with no well defined target body have received little attention. These Space Physics missions include some with extremely high energy requirements. Not only are these missions difficult, if not impractical, to accomplish with conventional chemical propulsion spacecraft, but they are also difficult to perform in the near future using projected electric-propulsion systems. Solar sail trajectories for two of the high energy Space Physics missions, a Solar Polar mission and a Heliopause mission, are examined in this paper. The object of the Solar Polar mission is to place 'a payload into a short period orbit around the Sun at a 90 degree inclination to either the ecliptic plane or the equatorial plane of the sun. A forerunner for this type of mission was the Ulysses spacecraft which used a gravity assist of Jupiter to place the spacecraft into a 90 degree inclination orbit around the Sun with a perihelion distance of around .55 AU. The orbital period for the Ulysses mission was around 5 years however, and future Solar Polar missions require many observational passes over the pole of the Sun each year thus implying a significantly shorter orbit period than that for Ulysses.

Sauer, Carl G., Jr.↗

NASA NTRS2000

Supplier's Status for Critical Solid Propellants, Explosive, and Pyrotechnic Ingredients

In the early 1970's a program was initiated at the Naval Surface Warfare Center/Indian Head Division (NSWC/IHDIV) to address the well-known problems associated with availability and suppliers of critical ingredients. These critical ingredients are necessary for preparation of solid propellants and explosives manufactured by the Navy. The objective of the program was to identify primary and secondary (or back-up) vendor information for these critical ingredients, and to develop suitable alternative materials if an ingredient is unavailable. In 1992 NSWC/IHDIV funded Chemical Propulsion Information Agency (CPIA) under a Technical Area Task (TAT) to expedite the task of creating a database listing critical ingredients used to manufacture Navy propellant and explosives based on known formulation quantities. Under this task CPIA provided employees that were 100 percent dedicated to the task of obtaining critical ingredient suppliers information, selecting the software and designing the interface between the computer program and the database users. TAT objectives included creating the Explosive Ingredients Source Database (EISD) for Propellant, Explosive and Pyrotechnic (PEP) critical elements. The goal was to create a readily accessible database, to provide users a quick-view summary of critical ingredient supplier's information and create a centralized archive that CPIA would update and distribute. EISD funding ended in 1996. At that time, the database entries included 53 formulations and 108 critical used to manufacture Navy propellant and explosives. CPIA turned the database tasking back over to NSWC/IHDIV to maintain and distribute at their discretion. Due to significant interest in propellant/explosives critical ingredients suppliers' status, the Propellant Development and Characterization Subcommittee (PDCS) approached the JANNAF Executive committee (EC) for authorization to continue the critical ingredient database work. In 1999, JANNAF EC approved the PDCS panel task. This paper is designed to emphasize the necessity of maintaining a JANNAF community supported database, which monitors PEP critical ingredient suppliers' status. The final product of this task is a user friendly, searchable database that provides a quick-view summary of critical ingredient supplier's information. This database must be designed to serve the needs of JANNAF and the propellant and energetic commercial manufacturing community as well. This paper provides a summary of the type of information to archive each critical ingredient.

Sims, B. L.↗

NASA NTRS2000

A Novel Approach to Exploring the Mars Polar Caps

The Martian polar caps contain some of the most important scientific sites on the planet. There is much interest in exploring them with a view to understanding their role in the Mars climate system. By gaining access to the stratigraphy of the polar terrain, it is probable that one can access the climate history of the planet. Additionally, investigations aimed at localizing subsurface water--liquid or solid--are not only of great scientific interest but are also germane to the long-term interests of the manned space flight program. A major difficulty with polar exploration is access. Current techniques using chemical propulsion, Holman transfers, and direct-entry landers with aeroshells have limited capability to access the polar terrain. For the near term the authors propose a new approach to solving this transportation issue by using Solar Electric Propulsion (SEP), recently flight demonstrated on NASA's DS1 Mission to an asteroid and a comet. For a longer-term approach there are additional ways in which access to Mars, as well as other planets, can be significantly improved. These include the use of Chaos orbit theory to enable transportation between LaGrange points in the solar system, gossamer structures enabling very low-mass mobility, and advanced ascent vehicles. In this paper the authors describe how a 1000-kG payload can be transported to the surface of Mars and a polar sample obtained and returned to Earth in less than five years using SEP. A vision of how this approach can be integrated into a long-term Mars exploration strategy building toward the future is also discussed.

Brophy, John R.↗

NASA NTRS2002

Thermal Examination of an Orbiting Cryogenic Fuel Depot

For many years NASA has been interested in the storage and transfer of cryogenic fuels in space. Lunar, L2 and other chemical propulsive space vehicle missions now have staged refueling needs that a fuel depot would satisfy. The depot considered is located in lower earth orbit. Many considerations must go into designing and building such a station. Multi-layer insulation systems, thermal shielding and low conductive structural supports are the principal means of protecting the system from excessive heat loss due to boiloff. This study focuses on the thermal losses associated with storing LH2 in a passively cooled fuel depot in a lower earth equatorial orbit. The corresponding examination looks at several configurations of the fuel depot. An analytical model has been developed to determine the thermal advantages and disadvantages of three different fuel depot configurations. Each of the systems consists of three Boeing rocket bodies arranged in various configurations. The first two configurations are gravity gradient stabilized while the third one is a spin-stabilized concept. Each concept was chosen for self-righting capabilities as well as the fuel settling capabilities, however the purpose of this paper is to prove which of the three concepts is the most efficient passively cooled system. The specific areas to be discussed are the heating time from the fusion temperature to the vaporization temperature and the amount of boiloff for a specific number of orbits. Each of the previous points is compared using various sun exposed surface areas of the tanks.

Hull, Patrick V.↗

NASA NTRS2002

Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE. While such a nozzle is a considerable idealization, it is clear that nozzle design and optimization will play a critical role in whether the performance potential of PDREs can be effectively realized in practice. In order to study PDRE nozzle issues with greater accuracy, a quasi-one-dimensional, finite-rate chemistry CFD code has been developed by the author. Comparisons of the code with both the previous MOC model and experimental data from Stanford University are reported. The effect of constant-gamma and finite-rate chemistry assumptions on the flowfield and performance is examined. Parametric studies of the effect of nozzle throat size and expansion ratio, at various blowdown pressure ratios, are reported.

Morris, Christopher I.↗

NASA NTRS2003

Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

Pulse detonation engines (PDB) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional gas turbines and rocket engines. The detonative mode of combustion employed by these devices offers a theoretical thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional engines. However, the unsteady blowdown process intrinsic to all pulse detonation devices has made realistic estimates of the actual propulsive performance of PDES problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models.

Morris, C. I.↗

NASA NTRS1960

An Analysis of the Corridor and Guidance Requirements for Supercircular Entry into Planetary Atmospheres

An analysis is presented of supercircular entry into a planet's atmosphere giving particular attention to the corridor through which spacecraft must be guided in order to accomplish various maneuvers. A dimensionless parameter based on conditions at the conic perigee altitude is introduced for characterizing supercircular entries and conveniently pre-scribing corridor widths associated with elliptic, parabolic, or hyperbolic approach trajectories. The analysis applies to vehicles of arbitrary weight, shape, and size. Illustrative calculations are made for Venus, Earth, Mars, Jupiter, and Titan. For nonlifting vehicles having fixed aerodynamic coefficients, curves are presented of dimensionless parameters from which can be calculated the maximum deceleration, maximum rate of laminar convective heating, and total laminar heat absorbed during single-pass entry at velocities up to twice circular velocity. For lifting vehicles, curves are presented of the maximum deceleration and overshoot boundary of an entry corridor; equations are presented for estimating laminar aerodynamic heating from the maximum deceleration. It is shown that the corridor width is independent of vehicle weight, dimensions, and drag coefficient, provided these are the same at the overshoot boundary as at undershoot. The corridors of certain planets can be broadened markedly by the application of aerodynamic lift; for example, the 10-earth-g corridor width for single- pass, nonlifting, parabolic entry is increased from 0 miles for Jupiter, 7 for Earth, and 8 for Venus, to 52, 51, and 52 miles, respectively, by employing a lift-drag ratio of 1. The use of aerodynamic lift does not increase appreciably the corridors of Mars and Titan. All corridor widths decrease rapidly as the entry velocity is increased. Terminal guidance requirements on accuracy of velocity and flight path angle for successfully entering various corridors are compared with analogous requirements for putting a satellite into orbit, for hitting the moon from the earth, and for achieving ICBM accuracy. Consideration is given to the terminal guidance problem involved in using a planet's atmosphere-rather than rocket fuel-to effect orbital transfers from heliocentric to planetocentric motion, thereby converting a hyperbolic approach trajectory to an elliptic orbit about the target planet. This fuel saving maneuver appears technologically feasible for certain planetary voyages, and implies the possibility of achieving a large reduction in required Earth lift-off weight of chemical propulsion systems.

Chapman, Dean R.↗