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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 55 records · Page 3
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NASA NTRS2004

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket, 1958-2002

During its maiden voyage in May 1962, a Centaur upper stage rocket, mated to an Atlas booster, exploded 54 seconds after launch, engulfing the rocket in a huge fireball. Investigation revealed that Centaur's light, stainless-steel tank had split open, spilling its liquid-hydrogen fuel down its sides, where the flame of the rocket exhaust immediately ignited it. Coming less than a year after President Kennedy had made landing human beings on the Moon a national priority, the loss of Centaur was regarded as a serious setback for the National Aeronautics and Space Administration (NASA). During the failure investigation, Homer Newell, Director of Space Sciences, ruefully declared: "Taming liquid hydrogen to the point where expensive operational space missions can be committed to it has turned out to be more difficult than anyone supposed at the outset." After this failure, Centaur critics, led by Wernher von Braun, mounted a campaign to cancel the program. In addition to the unknowns associated with liquid hydrogen, he objected to the unusual design of Centaur. Like the Atlas rocket, Centaur depended on pressure to keep its paper-thin, stainless-steel shell from collapsing. It was literally inflated with its propellants like a football or balloon and needed no internal structure to give it added strength and stability. The so-called "pressure-stabilized structure" of Centaur, coupled with the light weight of its high- energy cryogenic propellants, made Centaur lighter and more powerful than upper stages that used conventional fuel. But, the critics argued, it would never become the reliable rocket that the United States needed.

Dawson, Virginia P.↗
NASA NTRS2015

Solar Thermal Upper Stage Liquid Hydrogen Pressure Control Testing

High-energy cryogenic propellant is an essential element in future space exploration programs. Therefore, NASA and its industrial partners are committed to an advanced development/technology program that will broaden the experience base for the entire cryogenic fluid management community. Furthermore, the high cost of microgravity experiments has motivated NASA to establish government/aerospace industry teams to aggressively explore combinations of ground testing and analytical modeling to the greatest extent possible, thereby benefitting both industry and government entities. One such team consisting of ManTech SRS, Inc., Edwards Air Force Base, and Marshall Space Flight Center (MSFC) was formed to pursue a technology project designed to demonstrate technology readiness for an SRS liquid hydrogen (LH2) in-space propellant management concept. The subject testing was cooperatively performed June 21-30, 2000, through a partially reimbursable Space Act Agreement between SRS, MSFC, and the Air Force Research Laboratory. The joint statement of work used to guide the technical activity is presented in appendix A. The key elements of the SRS concept consisted of an LH2 storage and supply system that used all of the vented H2 for solar engine thrusting, accommodated pressure control without a thermodynamic vent system (TVS), and minimized or eliminated the need for a capillary liquid acquisition device (LAD). The strategy was to balance the LH2 storage tank pressure control requirements with the engine thrusting requirements to selectively provide either liquid or vapor H2 at a controlled rate to a solar thermal engine in the low-gravity environment of space operations. The overall test objective was to verify that the proposed concept could enable simultaneous control of LH2 tank pressure and feed system flow to the thruster without necessitating a TVS and a capillary LAD. The primary program objectives were designed to demonstrate technology readiness of the SRS concept at a system level as a first step toward actual flight vehicle demonstrations. More specific objectives included testing the pressure and feed control system concept hardware for functionality, operability, and performance. Valuable LH2 thermodynamic and fluid dynamics data were obtained for application to both the SRS concept and to future missions requiring space-based cryogen propellant management.

Moore, J. D.↗
NASA NTRS1962

A Survey and Evaluation of High Energy Liquid Chemical Propulsion Systems: Propellant Selection Criteria for Space Missions - Part 1

This report presents the results of a study to develop a procedure for evaluating liquid propellants in order (a) to select the most appropriate propellant (from among those under development) for each of several applications on each of the various missions in the NASA program, or (b) to select new propellants (from among those being proposed) for initiation or continuation of research and development. The analysis begins with a consideration of requirements--either for the specific application or for the various classes of applications. The known characteristics of the propellant or propellants to be evaluated are then put into a convenient form for evaluation. The next step is to determine whether or not there are requirements that simply cannot be met by the propellant. If the propellant passes this test, an optimum vehicle configuration using the propellant (and meeting all requirements) is estimated. (The configuration should be optimized with respect to the total resource consumption for all aspects of the mission, including R&D, production, logistics, and operation.) The total resource consumption for this configuration is then compared with that for similar configurations using other propellants (and meeting all requirements equally well). If all factors have been properly taken into account, this comparison of resource consumption will complete the evaluation. Such an evaluation may be performed several times, in increasing detail and with correspondingly increasing accuracy, as an R&D program proceeds, and the accuracy of the data as well as the cost of the next step in the program increase. The procedure is superior to those in common use in that it minimizes both the amount of analytical work and the number of points at which subjective value judgments are made.

LIQUID PROPELLANT↗
NASA NTRS1963

Apollo 2 Second Generation Lunar Exploration System Studies CFY 1963

The Apollo II Second Generation Lunar Exploration System includes the direct landing spacecraft which consists of cargo command module, service module, and landing module. The landing module is also capable of being used as the Lunar Landing Vehicle (LLV) for landing unmanned cargos consisting of shelter modules such as the Lunar Occupancy Payload and other cargo in support of lunar surface operations. High energy cryogenic propellants are utilized to permit direct landing, manned, or logistic missions with use of a single Saturn V class booster. In last year's studies, the LLV was configured for maximum payload and with consideration for the direct three-man landing and return mission. Light weight and low vehicle height above the lunar surface at touchdown were major objectives. Logistic cargos of more than 27,000 pounds landed on the Moon were achieved within the single Saturn V boost capability. For the manned mission, the lunar take-off weight was determined to be 28,000 pounds ready for the return-to-Earth portion of the mission. The command module utilized was an advanced light-weight design weighing 10,000 pounds including supporting subsystems. Cryogenic oxygen/hydrogen propulsion was again used for maximum propulsion efficiency. Study of the Lunar Occupancy Payload was also accomplished last year. This module was configured to serve as an early lunar shelter or outpost station or as a basic module of an integrated base module complex. Single and dual compartment versions, as well as special mission versions, were studied.

Matzenaur, J. O.↗
NASA NTRS1993

Advanced space propulsion concepts

The NASA Lewis Research Center has been actively involved in the evaluation and development of advanced spacecraft propulsion. Recent program elements have included high energy density propellants, electrode less plasma thruster concepts, and low power laser propulsion technology. A robust advanced technology program is necessary to develop new, cost-effective methods of spacecraft propulsion, and to continue to push the boundaries of human knowledge and technology.

Lapointe, Michael R.↗
NASA NTRS2006

Advanced Chemical Propulsion

Advanced Chemical Propulsion (ACP) provides near-term incremental improvements in propulsion system performance and/or cost. It is an evolutionary approach to technology development that produces useful products along the way to meet increasingly more demanding mission requirements while focusing on improving payload mass fraction to yield greater science capability. Current activities are focused on two areas: chemical propulsion component, subsystem, and manufacturing technologies that offer measurable system level benefits; and the evaluation of high-energy storable propellants with enhanced performance for in-space application. To prioritize candidate propulsion technology alternatives, a variety of propulsion/mission analyses and trades have been conducted for SMD missions to yield sufficient data for investment planning. They include: the Advanced Chemical Propulsion Assessment; an Advanced Chemical Propulsion System Model; a LOx-LH2 small pumps conceptual design; a space storables propellant study; a spacecraft cryogenic propulsion study; an advanced pressurization and mixture ratio control study; and a pump-fed vs. pressure-fed study.

Alexander, Leslie, Jr.↗
NASA NTRS1998

Launch Vehicle Performance with Solid Particle Feed Systems for Atomic Propellants

An analysis of launch vehicle Gross Liftoff Weight (GLOW) using high energy density atomic propellants with solid particle feed systems was conducted. The analyses covered several propellant combinations, including atoms of aluminum (Al), boron (B). carbon (C), and hydrogen (H) stored in a solid cryogenic particle, with a cryogenic liquid as the carrier fluid. Several different weight percents (wt%) for the liquid carrier were investigated and the gross lift off weight (GLOW) of the vehicles using the solid particle feed systems were compared with a conventional 02/H2 propellant vehicle. The potential benefits and effects of feed systems using solid particles in a liquid cryogenic fluid are discussed.

Palaszewski, Bryan↗
NASA NTRS2011

High energy, low temperature gelled bi-propellant formulation

The present invention is a bi-propellant system comprising a gelled liquid propane (GLP) fuel and a gelled MON-30 (70% N.sub.2O.sub.4+30% NO) oxidizer. The bi-propellant system is particularly well-suited for outer planet missions greater than 3 AU from the sun and also functions in earth and near earth environments. Additives such as powders of boron, carbon, lithium, and/or aluminum can be added to the fuel component to improve performance or enhance hypergolicity. The gelling agent can be silicon dioxide, clay, carbon, or organic or inorganic polymers. The bi-propellant system may be, but need not be, hypergolic.

Di Salvo, Roberto↗
NASA NTRS2013

High Energy, Low Temperature Gelled Bi-propellant Formulation Preparation Method

A method for preparing a gelled liquid propane (GLP) composition comprises the introduction of liquid propane into an evacuated mixing vessel containing a gellant and mixing the liquid propane with the gellant. A bi-propellant system comprising GLP is particularly well-suited for outer planet missions greater than 3 AU from the sun and also functions in earth and near earth environments. Additives such as powders of boron, carbon, lithium, and/or aluminum can be added improve performance or enhance hypergolicity. The gelling agent can be silicon dioxide, clay, carbon, or organic or inorganic polymers. The bi-propellant system may be, but need not be, hypergolic.

DiSalvo, Roberto↗
NASA NTRS1990

Comparison of super-high-energy-propulsion-systems based on metallic hydrogen propellant for ES to LEO space transportation

The application is studied of metallic H2 as a rocket propellant, which contains a specific energy of about 52 kcal/g in theory yielding a maximum specific impulse of 1700 s. With the convincing advantage of having a density 14 times that of conventional liquid H2/liquid O2 propellants, metallic H2 could satisfy the demands of advanced launch vehicle propulsion for the next millennium. Provided that there is an atomic metallic state of H2, and that this state is metastable at ambient pressure, which still is not proven, the results are given of the study of some important areas, which concern the production of metallic H2, the combustion, chamber cooling, and storage. The results show that the use of metallic H2 as rocket propellant could lead to revolutionary changes in space vehicle philosophy toward small size, small weight, and high performance single stage to orbit systems. The use of high metallic H2 mass fractions results in a dramatic reduction of required propellant volume, while gas temperatures in the combustion chamber exceed 5000 K. Furthermore, it follows, that H2 (liquid or slush) is the most favorable candidate as working fluid. Jet generated noise due to high exhaust velocities could be a problem.

Thierschmann, M.↗