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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Small rocket research and technology

Small chemical rockets are used on nearly all space missions. The small rocket program provides propulsion technology for civil and government space systems. Small rocket concepts are developed for systems which encompass reaction control for launch and orbit transfer systems, as well as on-board propulsion for large space systems and earth orbit and planetary spacecraft. Major roles for on-board propulsion include apogee kick, delta-V, de-orbit, drag makeup, final insertions, north-south stationkeeping, orbit change/trim, perigee kick, and reboost. The program encompasses efforts on earth-storable, space storable, and cryogenic propellants. The earth-storable propellants include nitrogen tetroxide (NTO) as an oxidizer with monomethylhydrazine (MMH) or anhydrous hydrazine (AH) as fuels. The space storable propellants include liquid oxygen (LOX) as an oxidizer with hydrazine or hydrocarbons such as liquid methane, ethane, and ethanol as fuels. Cryogenic propellants are LOX or gaseous oxygen (GOX) as oxidizers and liquid or gaseous hydrogen as fuels. Improved performance and lifetime for small chemical rockets are sought through the development of new predictive tools to understand the combustion and flow physics, the introduction of high temperature materials to eliminate fuel film cooling and its associated combustion inefficiency, and improved component designs to optimize performance. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Results indicate that modeling of the injector and combustion process in small rockets needs improvement. High temperature materials require the development of fabrication processes, a durability data base in both laboratory and rocket environments, and basic engineering property data such as strength, creep, fatigue, and work hardening properties at both room and elevated temperature. Promising materials under development include iridium-coated rhenium and a ceramic composite of mixed hafnium carbide and tantalum carbide reinforced with graphite fibers.

Schneider, Steven↗
NASA NTRS1993

Advanced materials for radiation-cooled rockets

The most common material system currently used for low thrust, radiation-cooled rockets is a niobium alloy (C-103) with a fused silica coating (R-512A or R-512E) for oxidation protection. However, significant amounts of fuel film cooling are usually required to keep the material below its maximum operating temperature of 1370 C, degrading engine performance. Also the R-512 coating is subject to cracking and eventual spalling after repeated thermal cycling. A new class of high-temperature, oxidation-resistant materials are being developed for radiation-cooled rockets, with the thermal margin to reduce or eliminate fuel film cooling, while still exceeding the life of silicide-coated niobium. Rhenium coated with iridium is the most developed of these high-temperature materials. Efforts are on-going to develop 22 N, 62 N, and 440 N engines composed of these materials for apogee insertion, attitude control, and other functions. There is also a complimentary NASA and industry effort to determine the life limiting mechanisms and characterize the thermomechanical properties of these materials. Other material systems are also being studied which may offer more thermal margin and/or oxidation resistance, such as hafnium carbide/tantalum carbide matrix composites and ceramic oxide-coated iridium/rhenium chambers.

Reed, Brian↗
NASA NTRS1994

Compact Simultaneous-beam Optical Strain Measurement System, Phase 5

Recent advances on the laser speckle strain measurement system under development at NASA Lewis Research Center have resulted in a compact, easy-to-use measurement package having many performance improvements over previous systems. NASA has developed this high performance optical strain measurement system for high temperature material testing applications. The system is based on I. Yamaguchi's two-beam speckle-shift strain measurement theory, and uses a new optical design that allows simultaneous recording of laser speckle patterns. This design greatly improves system response over previous implementations of the two-beam speckle-shift technique. The degree of immunity to transient rigid body motions is no longer dependent on the data transfer rate. The system automatically calculates surface strains at a frequency of about 5 Hz using a high speed digital signal processor in a personal computer. This system is fully automated, and can be operated remotely. This report describes the designs and methods used by the system, and shows low temperature strain test results obtained from small diameter tungsten-rhenium and palladium-chrome wires.

Lant, Christian T.↗
NASA NTRS1993

Advanced Small Rocket Chambers. Basic Program and Option 2: Fundamental Processes and Material Evaluation

Propellants, chamber materials, and processes for fabrication of small high performance radiation cooled liquid rocket engines were evaluated to determine candidates for eventual demonstration in flight-type thrusters. Both storable and cryogenic propellant systems were considered. The storable propellant systems chosen for further study were nitrogen tetroxide oxidizer with either hydrazine or monomethylhydrazine as fuel. The cryogenic propellants chosen were oxygen with either hydrogen or methane as fuel. Chamber material candidates were chemical vapor deposition (CVD) rhenium protected from oxidation by CVD iridium for the chamber hot section, and film cooled wrought platinum-rhodium or regeneratively cooled stainless steel for the front end section exposed to partially reacted propellants. Laser diagnostics of the combustion products near the hot chamber surface and measurements at the surface layer were performed in a collaborative program at Sandia National Laboratories, Livermore, CA. The Material Sample Test Apparatus, a laboratory system to simulate the combustion environment in terms of gas and material temperature, composition, and pressure up to 6 Atm, was developed for these studies. Rocket engine simulator studies were conducted to evaluate the materials under simulated combustor flow conditions, in the diagnostic test chamber. These tests used the exhaust species measurement system, a device developed to monitor optically species composition and concentration in the chamber and exhaust by emission and absorption measurements.

Jassowski, Donald M.↗
NASA NTRS1995

Long Life Testing of Oxide-Coated Iridium/Rhenium Rockets

22-N class rockets, composed of a rhenium (Re) substrate, an iridium (Ir) coating, and an additional composite coating consisting of Ir and a ceramic oxide, were tested on gaseous oxygen/gaseous hydrogen (GO2/GH2) propellants. Two rockets were tested, one for nearly 39 hours at a nominal mixture ratio (MR) of 4.6 and chamber pressure (Pc) of 469 kPa, and the other for over 13 hours at a nominal MR of 5.8 and 621 kPa Pc. Four additional Ir/Re rockets, with a composite Ir-oxide coating fabricated using a modified process, were also tested, including one for 1.3 hours at a nominal MR of 16.7 and Pc of 503 kPa. The long lifetimes demonstrated on low MR GO2/GH2 suggest greatly extended chamber lifetimes (tens of hours) in the relatively low oxidizing combustion environments of Earth storable propellants. The oxide coatings could also serve as a protective coating in the near injector region, where a still-mixing flowfield may cause degradation of the Ir layer. Operation at MR close to 17 suggests that oxide-coated Ir/Re rockets could be used in severely oxidizing combustion environments, such as high MR GO2/GH2, oxygen/hydrocarbon, and liquid gun propellants.

Reed, Brian D.↗
NASA NTRS1996

Brazing Dissimilar Metals

Brazing effective technique for joining ordinary structural metals to brittle, low-thermal-expansion refractory metals. Specifically, brazing process established for joining copper or nickel flanges to ends of vacuum-plasma-sprayed tungsten tubes and for joining stainless-steel flanges to ends of tubes made of alloy of molybdenum with 40 percent of rhenium.

Krotz, Phillip D.↗
NASA NTRS1997

A Rocket Engine for Mars Sample Return Using In Situ Propellants

Recent studies for the planned Mars sample return mission were reviewed and modified to utilize carbon monoxide and oxygen as potential in situ propellants. Based on these studies a representative full scale engine thrust of 2225 N (500 lbf) was selected as appropriate to demonstrate performance, and the design for that engine is presented. Previous experimental results combined with parametric analyses were used to define the geometry for the engine which operates on liquid carbon monoxide and liquid oxygen. The engine was constructed using a combination of high-temperature alloys and lightweight ceramics. The materials selected were hafnium oxide, iridium, rhenium, and carbon-carbon.

Linne, Diane L.↗
NASA NTRS1996

Analysis of Supercritical-Extracted Chelated Metal Ions From Mixed Organic-Inorganic Samples

Organic and inorganic contaminants of an environmental sample are analyzed by the same GC-MS instrument by adding an oxidizing agent to the sample to oxidize metal or metal compounds to form metal ions. The metal ions are converted to chelate complexes and the chelate complexes are extracted into a supercritical fluid such as CO2. The metal chelate extract after flowing through a restrictor tube is directly injected into the ionization chamber of a mass spectrometer, preferably containing a refractory metal filament such as rhenium to fragment the complex to release metal ions which are detected. This provides a fast, economical method for the analysis of metal contaminants in a sample and can be automated. An organic extract of the sample in conventional or supercritical fluid solvents can be detected in the same mass spectrometer, preferably after separation in a supercritical fluid chromatograph.

Sinha, Mahadeva P.↗
NASA NTRS1997

Evaluation of Heat Transfer and Thermal Stability of Supercritical JP-7 Fuel

A series of electrically heated tube experiments was conducted to investigate the potential of JP-7 as a coolant under conditions relevant to a Mach 8 propulsion system. The heat transfer capabilities, carbon deposition, and material compatibility of JP-7 at surface temperatures up to 1700 F (927 C) were tested in 0.125 in. diameter tubes of 304 SS, Inconel 617, Haynes 188, Haynes 230, and 50150 Moly-Rhenium. The heat transfer to the coolant was modeled well by a Dittus-Boelter correlation at lower heat fluxes. At higher heat fluxes, audible instabilities were observed and corresponded to a significant enhancement in the coolant heat transfer. The carbon deposition rates in these tests were comparable to those in previous experiments at lower heat fluxes and much longer residence times. This result suggests that alternative paths of the deposition mechanism may be enhanced under high heat flux test conditions. Microscopic investigation of the post test tubes indicated that there was a significant layer of ordered carbon deposits that had not been seen in the tests at lower heat flux.

Linne, Diane L.↗
NASA NTRS1998

Fabrication and Testing of Mo-Re Heat Pipes Embedded in Carbon/Carbon

Refractory-composite/heat-pipe-cooled wing an tail leading edges are being considered for use on hypersonic vehicles to limit maximum temperatures to values below material reuse limits and to eliminate the need to actively cool the leading edges. The development of a refractory-composite/heat-pipe-cooled leading edge has evolved from the design stage to the fabrication and testing of heat pipes embedded in carbon/carbon (C/C). A three-foot-long, molybdenum-rhenium heat pipe with a lithium working fluid was fabricated and tested at an operating temperature of 2460 F to verify the individual heat-pipe design. Following the fabrication of this heat pipe, three additional heat pipes were fabricated and embedded in C/C. The C/C heat-pipe test article was successfully tested using quartz lamps in a vacuum chamber in both a horizontal and vertical orientation. Start up and steady state data are presented for the C/C heat-pipe test article. Radiography and eddy current evaluations were performed on the test article.

Glass, David E.↗
NASA NTRS1998

Development Study of Cartridge/Crucible Tube Materials

The limitations of traditional alloys and the desire for improved performance for components is driving the increased utilization of refractory metals in tile space industry. From advanced propulsion systems to high temperature furnace components for microgravity processing, refractory metals are being used for their high melting temperatures and inherent chemical stability. Techniques have been developed to produce near net shape refractory metal components utilizing vacuum plasma spraying. Material utilization is very high, and laborious machining can be avoided. As-spray formed components have been tested and found to perform adequately. However, increased mechanical and thermal properties are needed. To improve these properties, post processing thermal treatments such as hydrogen sintering and vacuum annealing have been performed. Components formed from alloys of tungsten, rhenium, tantalum, niobium, and molybdenum are discussed and a metallurgical analyses detailing the results are presented. A qualitative comparison of mechanical properties is also included.

McKechnie, Timothy N.↗
NASA NTRS1998

On-Board Propulsion System Analysis of High Density Propellants

The impact of the performance and density of on-board propellants on science payload mass of Discovery Program class missions is evaluated. A propulsion system dry mass model, anchored on flight-weight system data from the Near Earth Asteroid Rendezvous mission is used. This model is used to evaluate the performance of liquid oxygen, hydrogen peroxide, hydroxylammonium nitrate, and oxygen difluoride oxidizers with hydrocarbon and metal hydride fuels. Results for the propellants evaluated indicate that the state-of-art, Earth Storable propellants with high performance rhenium engine technology in both the axial and attitude control systems has performance capabilities that can only be exceeded by liquid oxygen/hydrazine, liquid oxygen/diborane and oxygen difluoride/diborane propellant combinations. Potentially lower ground operations costs is the incentive for working with nontoxic propellant combinations.

Schneider, Steven J.↗
NASA NTRS1999

Vacuum Plasma Spray (VPS) Forming of Solar Thermal Propulsion Components Using Refractory Metals

The Thermal Spray Laboratory at NASA's Marshall Space Flight Center has developed and demonstrated a fabrication technique using Vacuum Plasma Spray (VPS) to form structural components from a tungsten/rhenium alloy. The components were assembled into an absorber cavity for a fully-functioning, ground test unit of a solar then-nal propulsion engine. The VPS process deposits refractory metal onto a graphite mandrel of the desired shape. The mandrel acts as a male mold, forming the required contour and dimensions of the inside surface of the deposit. Tungsten and tungsten/25% rhenium were used in the development and production of several absorber cavity components. These materials were selected for their high temperature (greater than 25000 C [greater than 4530 F]) strength. Each absorber cavity comprises 3 coaxial shells with two, double-helical flow passages through which the propellant gas flows. This paper describes the processing techniques, design considerations, and process development associated with forming these engine components.

Zimmerman, Frank R.↗
NASA NTRS1998

Vacuum Plasma Spray (VPS) Forming of Solar Thermal Propulsion Components Using Refractory Metals

The Thermal Spray Laboratory at NASA's Marshall Space Flight Center has developed and demonstrated a fabrication technique using Vacuum Plasma Spray (VPS) to form structural components from a tungsten/rhenium alloy. The components were assembled into an absorption cavity for a fully-functioning, ground test unit of a solar thermal propulsion engine. The VPS process deposits refractory metal onto a graphite mandrel of the desired shape. The mandrel acts as a male mold, forming the required contour and dimensions of the inside surface of the deposit. Tungsten and tungsten/25% rhenium were used in the development and production of several absorber cavity components. These materials were selected for their high temperature (less than 2500 C) strength. Each absorber cavity comprises 3 coaxial shells with two, double-helical flow passages through which the propellant gas flows. This paper describes the processing techniques, design considerations, and process development associated with forming these engine components.

Zimmerman, Frank↗
NASA NTRS1999

High Pressure Earth Storable Rocket Technology Program-Hipes Options 1/2 Report

Under the High Pressure Earth Storable Rocket Technology (HIPES) Program, TRW successfully completed testing of two 100 lbf thrust class rhenium chambers using N204-MMH. The first chamber was successfully fired for 4789 seconds of operating time with a maximum duration of 700 seconds. This chamber had been previously fired for 5230 seconds with N2O4-N2H4. The second chamber was successfully fired for 8085 seconds with a maximum firing duration of 1200 seconds. The Isp (specific impulse) for both chambers ranged from 323 lbf-sec/lbm to 330 lbf-sec/lbm.

Chazen, M. L.↗
NASA NTRS2000

Giving Bigger Satellites a Boost

Ultramet, Inc. has spurred a new process for producing rocket engine thrust chambers, through SBIR funding and the Glenn Research Center. High-temperature oxidation-resistant thruster materials are being produced in order to achieve high-temperature capability without sacrificing reliability. These thruster materials lead to an estimated three-percent improvement in propulsion system performance. To develop this material, Ultramet used a process called chemical vapor deposition (CVD). CVD involves heating precursors for metals, like iridium and rhenium, to temperatures at which they become gaseous. They are then deposited onto a mandrel, or spindle, layer-by-layer to produce high-density, highly resistant materials from the inside out.

Source record↗
NASA NTRS2003

Solar Thermal Propulsion Improvements at Marshall Space Flight Center

Solar Thermal Propulsion (STP) is a concept which operates by transferring solar energy to a propellant, which thermally expands through a nozzle. The specific impulse performance is about twice that of chemical combustions engines, since there is no need for an oxidizer. In orbit, an inflatable concentrator mirror captures sunlight and focuses it inside an engine absorber cavity/heat exchanger, which then heats the propellant. The primary application of STP is with upperstages taking payloads from low earth orbit to geosynchronous earth orbit or earth escape velocities. STP engines are made of high temperature materials since heat exchanger operation requires temperatures greater than 2500K. Refractory metals such as tungsten and rhenium have been examined. The materials must also be compatible with hot hydrogen propellant. MSFC has three different engine designs, made of different refractory metal materials ready to test. Future engines will be made of high temperature carbide materials, which can withstand temperatures greater than 3000K, hot hydrogen, and provide higher performance. A specific impulse greater than 1000 seconds greatly reduces the amount of required propellant. A special 1 OkW solar ground test facility was made at MSFC to test various STP engine designs. The heliostat mirror, with dual-axis gear drive, tracks and reflects sunlight to the 18 ft. diameter concentrator mirror. The concentrator then focuses sunlight through a vacuum chamber window to a small focal point inside the STP engine. The facility closely simulates how the STP engine would function in orbit. The flux intensity at the focal point is equivalent to the intensity at a distance of 7 solar radii from the sun.

Gerrish, Harold P.↗
NASA NTRS2004

Joint Development of a Fourth Generation Single Crystal Superalloy

A new, fourth generation, single crystal superalloy has been jointly developed by GE Aircraft Engines, Pratt & Whitney, and NASA. The focus of the effort was to develop a turbine airfoil alloy with long-term durability for use in the High Speed Civil Transport. In order to achieve adequate long-time strength improvements at moderate temperatures and retain good microstructural stability, it was necessary to make significant composition changes from 2nd and 3rd generation single crystal superalloys. These included lower chromium levels, higher cobalt and rhenium levels and the inclusion of a new alloying element, ruthenium. It was found that higher Co levels were beneficial to reducing both TCP precipitation and SRZ formation. Ruthenium caused the refractory elements to partition more strongly to the ' phase, which resulted in better overall alloy stability. The final alloy, EPM 102, had significant creep rupture and fatigue improvements over the baseline production alloys and had acceptable microstructural stability. The alloy is currently being engine tested and evaluated for advanced engine applications.

Walston, S.↗