Search NASA⌕ Search

SEARCH · Search NASA

Results for “rocket engine”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Mariner Venus/Mercury 1973 rocket engine assembly

The fabrication and test of rocket engine assemblies (REA) for Mariner Venus/Mercury 1973 are reported. The fabrication, assembly and flight acceptance test of seven REA's including the type approval test of one engine and fabrication of one additional kit consisting of detail parts for an engine ready for catalyst loading are presented. The MV/M '73 REA which is a nominal 51 lbs thrust monopropellant engine is described. Under steady state operation the specific impulse is not less than 228 lb-sec at 55 lb and 218.5 lb-sec at 10 lb thrust varying linearly between these limits. The characteristic velocity is not less than 4100 ft/sec at any thrust level.

Snoke, D. R.↗
NASA NTRS1974

Centrifugal pumps for rocket engines

The use of centrifugal pumps for rocket engines is described in terms of general requirements of operational and planned systems. Hydrodynamic and mechanical design considerations and techniques and test procedures are summarized. Some of the pump development experiences, in terms of both problems and solutions, are highlighted.

Campbell, W. E.↗
NASA NTRS1982

Reusable Rocket Engine Maintenance Study

Approximately 85,000 liquid rocket engine failure reports, obtained from 30 years of developing and delivering major pump feed engines, were reviewed and screened and reduced to 1771. These were categorized into 16 different failure modes. Failure propagation diagrams were established. The state of the art of engine condition monitoring for in-flight sensors and between flight inspection technology was determined. For the 16 failure modes, the potential measurands and diagnostic requirements were identified, assessed and ranked. Eight areas are identified requiring advanced technology development.

Macgregor, C. A.↗
NASA NTRS2002

Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE is described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of several different PDRE nozzle configurations is reported.

Morris, Christopher I.↗
NASA NTRS1991

Reusable rocket engine intelligent control system framework design, phase 2

Elements of an advanced functional framework for reusable rocket engine propulsion system control are presented for the Space Shuttle Main Engine (SSME) demonstration case. Functional elements of the baseline functional framework are defined in detail. The SSME failure modes are evaluated and specific failure modes identified for inclusion in the advanced functional framework diagnostic system. Active control of the SSME start transient is investigated, leading to the identification of a promising approach to mitigating start transient excursions. Key elements of the functional framework are simulated and demonstration cases are provided. Finally, the advanced function framework for control of reusable rocket engines is presented.

Nemeth, ED↗
NASA NTRS2003

Pulse Detonation Rocket Engine Research at NASA Marshall

This viewgraph representation provides an overview of research being conducted on Pulse Detonation Rocket Engines (PDRE) by the Propulsion Research Center (PRC) at the Marshall Space Flight Center. PDREs have a theoretical thermodynamic advantage over Steady-State Rocket Engines (SSREs) although unsteady blowdown processes complicate effective use of this advantage in practice; PRE is engaged in a fundamental study of PDRE gas dynamics to improve understanding of performance issues. Topics covered include: simplified PDRE cycle, comparison of PDRE and SSRE performance, numerical modeling of quasi 1-D rocket flows, time-accurate thrust calculations, finite-rate chemistry effects in nozzles, effect of F-R chemistry on specific impulse, effect of F-R chemistry on exit species mole fractions and PDRE performance optimization studies.

Morris, Christopher I.↗
NASA NTRS1999

Catalyst Development for Hydrogen Peroxide Rocket Engines

The development of various catalysts of hydrogen peroxide was conducted for the applications of liquid rocket engines. The catalyst development includes silver screen technology, solid catalyst technology, and homogeneous catalyst technology. The silver screen technology development was performed with 85% (by weight) hydrogen peroxide. The results of this investigation were used as the basis for the catalyst design of a pressure-fed liquid-fueled upper stage engine. Both silver-plated nickel 200 screens and pure silver screens were used as the active metal catalyst during the investigation, The data indicate that a high decomposition efficiency (greater than 90%) of 85% hydrogen peroxide can be achieved at a bed loading of 0.5 lbm/sq in/sec with both pure silver and silver plated screens. Samarium oxide coating, however, was found to retard the decomposition process and the catalyst bed was flooded at lower bed loading. A throughput of 200 lbm of hydrogen peroxide (1000 second run time) was tested to evaluate the catalyst aging issue and performance degradation was observed starting at approximately 400 seconds. Catalyst beds of 3.5 inch in diameter was fabricated using the same configuration for a 1,000-lbf rocket engine. High decomposition efficiency was obtained with a low pressure drop across the bed. Solid catalyst using precious metal was also developed for the decomposition of hydrogen peroxide from 85% to 98% by weight. Preliminary results show that the catalyst has a strong reactivity even after 15 minutes of peroxide decomposition. The development effort also includes the homogeneous catalyst technology. Various non-toxic catalysts were evaluated with 98% peroxide and hydrocarbon fuels. The results of open cup drop tests indicate an ignition delay around 11 ms.

Morlan, P. W.↗
NASA NTRS2024

ECI Final Report: Closing of Critical Technology Gaps for Rotating Detonation Rocket Engines

NASA has successfully fired several configurations of a dual regenerative 10K lbf class Rotating Detonation Rocket Engines using cryogenic liquid methane, liquid hydrogen, and RP-1 as fuel and liquid or gaseous oxygen as the oxidizer. In addition, 3 hot fire test phases were conducted to evaluate parametrically varying combustor geometries using a subscale 500-1K lbf class thruster RDRE. Several studies were successfully completed to evaluate the impacts of contraction ratio, annulus L’, heat transfer through calorimetry, and varying injector designs. Laser powder bed fusion GRCop-42, GRX-810, and C-103 alloys were used for the outer bodies, inner bodies, nozzles, and injector variations. The development of RDRE technology is only now possible with additive manufacturing techniques and these specialized alloys. Critical design parameters, design rules of thumb, and operability limitations have been identified. Major lessons learned in the development of critical components have also been documented. All critical technology gaps have been demonstrated at one scale or another and closed. A major finding of this work is that the combustion efficiency of the RDRE is far superior to that of the traditional liquid rocket engine requiring an order of magnitude lower chamber volume and residence time to react. Because of this, the design trade space opens up significantly. More compact chamber geometries are now possible with potential mass savings and significant length savings possible to flight geometries. In addition, the ceiling for the technology, in terms of Isp, has not yet been reached. It is likely that an additional ~10% better Isp over equivalent state-of-the-art combustion devices may be achieved over the next decade should the technology continue to be developed. This technology has been shown to be viable for near term infusion (2-5 years) into space missions and enable cost effective space access for US industry. This final report documents all major findings from this early career initiative (ECI) project and lays out what is needed post ECI.

Thomas Teasley↗
NASA NTRS1991

Fiber optics in liquid propellant rocket engine environments

Fiber optics have recently been seen to offer several major benefits in liquid-fuel rocket engine applications. Fiber-optic sensors can provide measurements that cannot be made with conventional techniques. Fiber optics also can reduce harness weight, provide lightning immunity, and increase frequency response. This paper discusses the results of feasibility testing optical fibers in simulated liquid-fuel rocket engine environments. The environments included cryogenic and high temperatures, and high vibration levels.

Delcher, R.↗
NASA NTRS2021

Carbon- and Ceramic-Matrix Composite Nozzle Extensions for Liquid Rocket Engines

Carbon-carbon (C-C) composite and ceramic matrix composite (CMC) nozzle extensions are being considered and developed for use with a variety of cryogenic liquid rocket engine systems. Some of the potential applications include in-space, lunar lander, and launch vehicle upper stage propulsion systems. Composite material systems of interest include C-C and carbon/silicon-carbide (C-SiC) composites, as well as modified versions of such composites (coatings, mixed matrices, oxidation inhibitors, etc.). Nozzle extension materials and fabrication methods under development by industry support both National Aeronautics and Space Administration (NASA) and Department of Defense (DOD) requirements, as well as those of the broader Commercial Space industry. Marshall Space Flight Center (MSFC) efforts are aimed at both (a) assisting industry with the development of composite nozzle extensions for cryogenic liquid rocket engines, and (b) developing and demonstrating capabilities for characterizing, testing, and qualifying such nozzle extensions. Results of recent MSFC nozzle-extension characterization efforts and engine firing demonstrations will be presented.

carbon-carbon↗
NASA NTRS1991

Ceramic composites for rocket engine turbines

The use of ceramic materials in the hot section of the fuel turbopump of advanced reusable rocket engines promises increased performance and payload capability, improved component life and economics, and greater design flexibility. Severe thermal transients present during operation of the Space Shuttle Main Engine (SSME), push metallic components to the limit of their capabilities. Future engine requirements might be even more severe. In phase one of this two-phase program, performance benefits were quantified and continuous fiber reinforced ceramic matrix composite components demonstrated a potential to survive the hostile environment of an advanced rocket engine turbopump.

Herbell, Thomas P.↗
NASA NTRS1991

Ceramic composites for rocket engine turbines

The use of ceramic materials in the hot section of the fuel turbopump of advanced reusable rocket engines promises increased performance and payload capability, improved component life and economics, and greater design flexibility. Severe thermal transients present during operation of the Space Shuttle Main Engine (SSME), push metallic components to the limit of their capabilities. Future engine requirements might be even more severe. In phase one of this two-phase program, performance benefits were quantified and continuous fiber reinforced ceramic matrix composite components demonstrated a potential to survive the hostile environment of an advaced rocket engine turbopump.

Herbell, Thomas P.↗
NASA NTRS2005

Towards Rocket Engine Components with Increased Strength and Robust Operating Characteristics

High-energy rotating machines, powering liquid propellant rocket engines, are subject to various sources of high and low cycle fatigue generated by unsteady flow phenomena. Given the tremendous need for reliability in a sustainable space exploration program, a fundamental change in the design methodology for engine components is required for both launch and space based systems. A design optimization system based on neural-networks has been applied and demonstrated in the redesign of the Space Shuttle Main Engine (SSME) Low Pressure Oxidizer Turbo Pump (LPOTP) turbine nozzle. One objective of the redesign effort was to increase airfoil thickness and thus increase its strength while at the same time detuning the vane natural frequency modes from the vortex shedding frequency. The second objective was to reduce the vortex shedding amplitude. The third objective was to maintain this low shedding amplitude even in the presence of large manufacturing tolerances. All of these objectives were achieved without generating any detrimental effects on the downstream flow through the turbine, and without introducing any penalty in performance. The airfoil redesign and preliminary assessment was performed in the Exploration Technology Directorate at NASA ARC. Boeing/Rocketdyne and NASA MSFC independently performed final CFD assessments of the design. Four different CFD codes were used in this process. They include WIL DCA T/CORSAIR (NASA), FLUENT (commercial), TIDAL (Boeing Rocketdyne) and, a new family (AardvarWPhantom) of CFD analysis codes developed at NASA MSFC employing LOX fluid properties and a Generalized Equation Set formulation. Extensive aerodynamic performance analysis and stress analysis carried out at Boeing Rocketdyne and NASA MSFC indicate that the redesign objectives have been fully met. The paper presents the results of the assessment analysis and discusses the future potential of robust optimal design for rocket engine components.

Marcu, Bogdan↗