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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 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Evaluation of Propulsion Systems for Mach 2.4 and 3.2 High Speed Civil Transports

Advanced technology high speed civil transport (HSCT) engine cycles equipped with low emissions combustors and low noise exhaust nozzles were investigated. The objectives were as follows: (1) evaluate a range of cycle variables for engine types appropriate for Mach 2.4 and 3.2 HSCT; (2) select the most promising propulsion systems; and (3) identify and define the critical/key technology development and demonstration programs needed to be conducted prior to initiation of the product engine development and certification program. Two types of variable engines were studied: turbine bypass engines (TBE) and variable stream control engines (VSCE). This study identified payoffs in terms of aircraft gross weight, airline economics, stratospheric cruise emissions, and airport noise improvements. Comparison between the the TBE and VSCE showed the TBE to be clearly superior for Mach 2.4 and 3.2 cruise cruise operation. Critical technology areas were also identified which included the following: (1) Low emissions combustor; (2) Low noise exhaust system; (3) High temperature materials and coatings; (4) Advanced turbine cooling and aerodynamics; (5) Integrated propulsion system (inlet/engine/nozzle) and airframe controls; (6) Advanced fan/compressor aerodynamics; and (7) Magnetic bearings.

Turbine Bypass Engine (TBE)↗
NASA NTRS2002

Potential Operating Orbits for Fission Electric Propulsion Systems Driven by the SAFE-400

Safety must be ensured during all phases of space fission system design, development, fabrication, launch, operation, and shutdown. One potential space fission system application is fission electric propulsion (FEP), in which fission energy is converted into electricity and used to power high efficiency (Isp greater than 3000s) electric thrusters. For these types of systems it is important to determine which operational scenarios ensure safety while allowing maximum mission performance and flexibility. Space fission systems are essentially nonradioactive at launch, prior to extended operation at high power. Once high power operation begins, system radiological inventory steadily increases as fission products build up. For a given fission product isotope, the maximum radiological inventory is typically achieved once the system has operated for a length of time equivalent to several half-lives. After that time, the isotope decays at the same rate it is produced, and no further inventory builds in. For an FEP mission beginning in Earth orbit, altitude and orbital lifetime increase as the propulsion system operates. Two simultaneous effects of fission propulsion system operation are thus (1) increasing fission product inventory and (2) increasing orbital lifetime. Phrased differently, as fission products build up, more time is required for the fission products to naturally convert back into non-radioactive isotopes. Simultaneously, as fission products build up, orbital lifetime increases, providing more time for the fission products to naturally convert back into non-radioactive isotopes. Operational constraints required to ensure safety can thus be quantified.

Houts, Mike↗
NASA NTRS1973

Propulsion system for research VTOL transports.

In anticipation of an eventual VTOL requirement for civil aviation, NASA has been conducting studies directed toward determining and developing the technology required for a commercial VTOL transport. The commercial transport configurations of such a VTOL transport are briefly reviewed, the propulsion system specifications and components developed by the engine study contractor are presented and described, and methods for using the lift-propulsion system for aircraft attitude control are discussed.

Gertsma, L. W.↗
NASA NTRS2006

Status of a Power Processor for the Prometheus-1 Electric Propulsion System

NASA is developing technologies for nuclear electric propulsion for proposed deep space missions in support of the Exploration initiative under Project Prometheus. Electrical power produced by the combination of a fission-based power source and a Brayton power conversion and distribution system is used by a high specific impulse ion propulsion system to propel the spaceship. The ion propulsion system include the thruster, power processor and propellant feed system. A power processor technology development effort was initiated under Project Prometheus to develop high performance and lightweight power-processing technologies suitable for the application. This effort faces multiple challenges including developing radiation hardened power modules and converters with very high power capability and efficiency to minimize the impact on the power conversion and distribution system as well as the heat rejection system. This paper documents the design and test results of the first version of the beam supply, the design of a second version of the beam supply and the design and test results of the ancillary supplies.

Pinero, Luis R.↗
NASA NTRS1980

An electric vehicle propulsion system's impact on battery performance: An overview

The performance of two types of batteries, lead-acid and nickel-zinc, was measured as a function of the charging and discharging demands anticipated from electric vehicle propulsion systems. The benefits of rapid high current charging were mixed: although it allowed quick charges, the energy efficiency was reduced. For low power (overnight) charging the current wave shapes delivered by the charger to the battery tended to have no effect on the battery cycle life. The use of chopper speed controllers with series traction motors resulted in a significant reduction in the energy available from a battery whenever the motor operates at part load. The demand placed on a battery by an electric vehicle propulsion system containing electrical regenerative braking confirmed significant improvment in short term performance of the battery.

Bozek, J. M.↗
NASA NTRS2019

Ice-Crystal Icing Accretion Studies at the NASA Propulsion Systems Laboratory

This paper describes an ice-crystal icing experiment conducted at the NASA Propulsion System Laboratory during June 2018. This test produced ice shape data on an airfoil for different test conditions similar to those inside the compressor region of a turbo-fan jet engine. Mixed-phase icing conditions were generated by partially freezing out a water spray using the relative humidity of flow as the primary parameter to control freeze-out. The paper presents the ice shape data and associated conditions which include pressure, velocity, temperature, humidity, total water content, melt ratio, and particle size distribution. The test featured a new instrument traversing system which allowed surveys of the flow and cloud. The purpose of this work was to provide experimental ice shape data and associated conditions to help develop and validate ice-crystal icing accretion models. The results support previous experimental observations of a minimum melt-ratio threshold for accretion to occur as well as the existence of a plateau region where the icing severity is high for a range of melt ratios. However, a maximum limit for melt ratio, which is suggested in the ice crystal icing literature, was not observed perhaps complicated by the potential for some supercooling of the water at these conditions.

Engine Icing↗
NASA NTRS1991

Numerical propulsion system simulation: An interdisciplinary approach

The tremendous progress being made in computational engineering and the rapid growth in computing power that is resulting from parallel processing now make it feasible to consider the use of computer simulations to gain insights into the complex interactions in aerospace propulsion systems and to evaluate new concepts early in the design process before a commitment to hardware is made. Described here is a NASA initiative to develop a Numerical Propulsion System Simulation (NPSS) capability.

Nichols, Lester D.↗
NASA NTRS1991

Numerical propulsion system simulation - An interdisciplinary approach

The tremendous progress being made in computational engineering and the rapid growth in computing power that is resulting from parallel processing now make it feasible to consider the use of computer simulations to gain insights into the complex interactions in aerospace propulsion systems and to evaluate new concepts early in the design process before a commitment to hardware is made. Described here is a NASA initiative to develop a Numerical Propulsion System Simulation (NPSS) capability.

Nichols, Lester D.↗
NASA NTRS1993

Spacecraft propulsion systems test capability at the NASA White Sands Test Facility

The NASA White Sands Facility (WSTF), a component insallation of the Johnson Space Center, is located on a 94-square-mile site in southwestern New Mexico. WSTF maintains many unique capabilities to support its mission to test and evaluate spacecraft materials, components, and propulsion systems to enable the safe human exploration and utilization of space. WSTF has tested over 340 rocket engines with more than 2.5 million firings to date. Included are propulsion system testing for Apollo, Shuttle, and now Space Station as well as unmanned spacecraft such as Viking, Pioneer, and Mars Observer. This paper describes the current WSTF propulsion test facilities and capabilities.

Baker, Pleddie↗
NASA NTRS1979

Preflight status of the External Tank portion of the Space Shuttle Main Propulsion System

The development progress of the External Tank (ET) of the Shuttle Main Propulsion System (MPS) is described. 87% of ET propulsion components have satisfactorily completed qualification tests in preparation for the first Shuttle flight and their performance has been verified in the Main Propulsion Test Article (MPTA) firings of up to 100 seconds duration. The design changes in the LH2 tank pressurant diffuser, the LO2 feed system, and the ground test LO2 tank are discussed. Propellant loading information from MPTA tests may be used to predict loading operations and performance at KSC and VAFB, while ullage pressure profiles and boil-off rates during loading, chilldown rates, and final bulk density data have been obtained with geyser suppression evaluated and loading procedures established and verified. The MPTA test experience provided data to evaluate pressurization system and ET to Orbiter interface propellant performance noting thus that the fluid interface requirements will be satisfied for the first Shuttle flight. The Intertank inerting system was found to be effective in minimizing hazardous gas mixtures and full scale separation tests of the ET to Ground Umbilical Carrier Plate system were successfully completed.

Norquist, L.↗
NASA NTRS1998

Propellant Feed Subsystem for the X-34 Main Propulsion System

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major NMS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. This paper reviews major design and analysis aspects of the X-34 propellant feed subsystem of the X-34 Main Propulsion System. Topics include system requirements, system design, the integration of flight and feed system performance, propellant acquisition at engine start, and propellant tank terminal drain.

McDonald, J. P.↗
NASA NTRS2014

Design, Integration, Certification and Testing of the Orion Crew Module Propulsion System

The Orion Multipurpose Crew Vehicle (MPCV) is NASA's next generation spacecraft for human exploration of deep space. Lockheed Martin is the prime contractor for the design, development, qualification and integration of the vehicle. A key component of the Orion Crew Module (CM) is the Propulsion Reaction Control System, a high‐flow hydrazine system used during re‐entry to orient the vehicle for landing. The system consists of a completely redundant helium (GHe) pressurization system and hydrazine fuel system with monopropellant thrusters. The propulsion system has been designed, integrated, and qualification tested in support of the Orion program's first orbital flight test, Exploration Flight Test One (EFT‐1), scheduled for 2014. A subset of the development challenges and lessons learned from this first flight test campaign will be discussed in this paper for consideration when designing future spacecraft propulsion systems. The CONOPS and human rating requirements of the CM propulsion system are unique when compared with a typical satellite propulsion reaction control system. The system requires a high maximum fuel flow rate. It must operate at both vacuum and sea level atmospheric pressure conditions. In order to meet Orion's human rating requirements, multiple parts of the system must be redundant, and capable of functioning after spacecraft system fault events.

McKay, Heather↗
NASA NTRS2010

2010 JPC Abstract: Ares I First Stage Propulsion System Status

In November 2005, NASA created the Constellation Program to develop an entirely new fleet of spacecraft to include the Ares I Crew Launch Vehicle and Ares V Cargo Launch vehicles. This mission architecture included the Orion capsule (which would be used to transport astronauts to low-Earth orbit and beyond), the Altair lunar lander, and an Earth departure stage. The Ares First Stage Team has made significant progress on the design of a propulsion system to meet the objectives of the Constellation Program. Work on a first stage element propulsion system capable of lofting a new fleet of spacecraft is well underway. To minimize technical risks and development costs, the Solid Rocket Boosters (SRBs) of Shuttle served as a starting point in the design of a new motor that would meet the requirements of those new vehicles. This new propulsive element will provide greater total impulse utilizing a fifth segment to loft a safer, more powerful fleet of space flight vehicles. Performance requirements, basic architecture, and obsolescence issues were all factors in determining the new first stage element design and configuration. Early efforts focused on creating designs that would be capable of supporting the requisite loads and environments. While the motor casings are Shuttle legacy, because of Ares I s unique in-line configuration, the first stage will require entirely new forward structures (forward skirt, forward skirt extension, aeroshell, and frustum) and a modified systems tunnel. The use of composites facilitated a change in the geometry, which in turn afforded the ability to focus strength where it was needed without additional mass. The Ares First Stage rocket motor casting tooling was designed and built to achieve a propellant grain geometry that produces the specific required ballistic profile. The new propellant formulation is a polybutadiene acrylonitrile (PBAN) copolymer, which has been modified to attain the desired burn rate and retain adequate tailoring capability.

Priskos, Alex S.↗
NASA NTRS1999

HSR Propulsion System Studies: A Status Report on the Down-Select Process

When the HSR program began there was widespread belief that a simple and familiar turbojet-like engine coupled to an advanced technology mixer-ejector nozzle was the propulsion system of choice for achieving FAR 36-Stage 3 noise requirements. Our ability to quickly demonstrate a practical 20(+) dB suppression nozzle was confidently presumed by many. Our rate of progress towards that objective, however, has been somewhat humbling. At the moment we are reasonably confident of achieving about 15 dB suppression with a mixer-ejector nozzle designed for a high specific thrust turbojet-like cycle. Therefore, if we make no further suppression progress and conservatively assume no new operational procedures such as programmed lapse rate (PLR), then meeting the Stage 3 goal requires a large amount of engine and/or wing oversizing which is economically prohibitive. The scenario is further aggravated by the possibility of eventually needing to comply with even more stringent regulations (Stage 4). While this status may be somewhat disappointing to some, it must be remembered that the HSR program plan involves two generations of mixer-ejector nozzles beyond the current generation I nozzle designs. It is premature to conclude that we cannot design a practical 20(+) dB mixer-ejector nozzle. On the other hand, it is prudent to consider alternative solutions to the noise problem. Thus, we are investigating four other propulsion system concepts.

Strack, W. C.↗
NASA NTRS2008

Summary of Propulsion System Needs in Support of Project Constellation

In January 2004, the President of the United States established the Vision for Space Exploration (VSE) to return man to the moon and ultimately to extend manned space travel to Mars. This paper will summarize the manned space flight liquid propulsion system needs in support of Project Constellation over the next 10 years. It will include all engine needs to return man to the moon. An overview of engines currently under contract, those baselined but not yet under contract, and those engine needs that hav.e yet to be initiated. Project Constellation includes the components as shown Figure 1. Liquid propulsion systems supporting the manned portion of these elements include the following: the Crew Exploration Vehicle named Orion (crew module reaction control system (CMRCS), service module Orion Main Engine (OME), service module auxiliary RCS, and service module reaction control system (SMRCS)), the Crew Launch Vehicle named Ares 1 (J2X upper stage, first stage roll control system, second stage reaction control system, and the Ares I-X roll control system), the Heavy Lift Launch Vehicle named Ares V (RS68B first stage booster, J-2X upper stage, roll control systems, and the Earth Departure Stage (EDS) (powered by the same Ares V Upper Stage J-2X), and the Lunar Lander named Altair with both descent and ascent stages (lunar orbit insertion and descent main engine, ascent main engine, and attitude control systems for both stages). In addition, there may be additional engine needs for early demonstrators, but those will not be speculated on as part of this paper. Also, other portions of the VSE architecture, including the planned Orion abort demonstrations and the Lunar Precursor Robotic Program, are not addressed here as they either use solid motors or are focused on unmanned precursor missions.

Sumrall, Phil↗
NASA NTRS2017

In-Space Chemical Propulsion Systems Roadmap

In-space propulsion begins where the launch vehicle upper stage leaves off, performing the functions of primary propulsion, reaction control, station keeping, precision pointing, and orbital maneuvering. The main engines used in space provide the primary propulsive force for orbit transfer, planetary trajectories and extra planetary landing and ascent. The reaction control and orbital maneuvering systems provide the propulsive force for orbit maintenance, position control, station keeping, and spacecraft attitude control. Advanced in-space propulsion technologies will enable much more effective exploration of our Solar System and will permit mission designers to plan missions to "fly anytime, anywhere, and complete a host of science objectives at the destinations" with greater reliability and safety. With wide range of possible missions and candidate propulsion technologies, the question of which technologies are "best" for future missions is a difficult one. A portfolio of propulsion technologies should be developed to provide optimum solutions for a diverse set of missions and destinations. A large fraction of the rocket engines in use today are chemical rockets; that is, they obtain the energy needed to generate thrust by chemical reactions to create a hot gas that is expanded to produce thrust. A significant limitation of chemical propulsion is that it has a relatively low specific impulse (Is, or thrust per mass flow rate of propellant). A significant improvement (>30%) in Is can be obtained by using cryogenic propellants, such as liquid oxygen and liquid hydrogen, for example. Historically, these propellants have not been applied beyond upper stages.

Oxidizers↗