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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Niobium-aluminum base alloys having improved, high temperature oxidation resistance

A niobium-aluminum base alloy having improved oxidation resistance at high temperatures and consisting essentially of 48%-52% niobium, 36%-42% aluminum, 4%-10% chromium, 0%-2%, more preferably 1%-2%, silicon and/or tungsten with tungsten being preferred, and 0.1%-2.0% of a rare earth selected from the group consisting of yttrium, ytterbium and erbium. Parabolic oxidation rates, k.sub.p, at 1200.degree. C. range from about 0.006 to 0.032 (mg/cm.sup.2).sup.2 /hr. The new alloys also exhibit excellent cyclic oxidation resistance.

Hebsur, Mohan G.↗
NASA NTRS2019

GRCop-42 Development and Hot-fire Testing Using Additive Manufacturing Powder Bed Fusion for Channel-Cooled Combustion Chambers

GRCop-42 is a high conductivity, high-strength dispersion strengthened copper-alloy for use in high heat flux applications such as liquid rocket engine combustion devices. This alloy is part of the family of NASA-developed GRCop, copper-chrome-niobium alloys. GRCop alloys were developed for harsh environments specific to regeneratively-cooled combustion chambers and nozzles with good oxidation resistance. Significant development was completed on the GRCop-84 and GRCop-42 alloys in the extruded wrought form demonstrating feasibility for combustion chambers. NASA has recently developed a process for additive manufacturing, specifically Powder Bed Fusion (PBF) or Selective Laser Melting (SLM), of GRCop-42 to establish parameters, characterize the material, and complete testing of components with complex internal features. This evolution of the GRCop-42 was based on the successful predecessor development work using GRCop-84 with the motivation of establishing a new copper-alloy option for use in NASA, government, and industry programs with SLM. A few advantages have been shown with the GRCop-42 that include higher conductivity and faster build speeds over the GRCop-84, and a simplified powder supply chain. Initial property development has shown that it is possible to produce high density builds with strengths equivalent to wrought GRCop-42 and a conductivity greater than GRCop-84. The GRCop-42 has completed process development and initial properties have been established. Several demonstrator combustion chambers have also been fabricated with the SLM GRCop-42 that include integral channels and closeouts. Additional test units have been fabricated and are completing substantial hot-fire testing to demonstrate performance of the material, process, and design.

Gradl, Paul R.↗
NASA NTRS1991

Advanced materials for space nuclear power systems

The overall philosophy of the research was to develop and characterize new high temperature power conversion and radiator materials and to provide spacecraft designers with material selection options and design information. Research on three candidate materials (carbide strengthened niobium alloy PWC-11 for fuel cladding, graphite fiber reinforced copper matrix composites for heat rejection fins, and tungsten fiber reinforced niobium matrix composites for fuel containment and structural supports considered for space power system applications is discussed. Each of these types of materials offers unique advantages for space power applications.

Titran, Robert H.↗
NASA NTRS1991

Advanced materials for space nuclear power systems

The overall philosophy of the research was to develop and characterize new high temperature power conversion and radiator materials and to provide spacecraft designers with material selection options and design information. Research on three candidate materials (carbide strengthened niobium alloy PWC-11 for fuel cladding, graphite fiber reinforced copper matrix composites for heat rejection fins, and tungsten fiber reinforced niobium matrix composites for fuel containment and structural supports) considered for space power system applications is discussed. Each of these types of materials offers unique advantages for space power applications.

Titran, Robert H.↗
NASA NTRS1996

Characterization of SiC Fiber (SCS-6) Reinforced-Reaction-Formed Silicon Carbide Matrix Composites

Silicon carbide fiber (SCS-6) reinforced-reaction-formed silicon carbide matrix composites were fabricated using a reaction-forming process. Silicon-2 at.% niobium alloy was used as an infiltrant instead of pure silicon to reduce the amount of free silicon in the matrix after reaction forming. The matrix primarily consists of silicon carbide with a bimodal grain size distribution. Minority phases dispersed within the matrix are niobium disilicide (NbSi2), carbon, and silicon. Fiber pushout tests on these composites determined a debond stress of approximately 67 MPa and a frictional stress of approximately 60 MPa. A typical four-point flexural strength of the composite is 297 MPa (43.1 KSi). This composite shows tough behavior through fiber pullout.

Singh, M.↗
NASA NTRS1995

Characterization of SiC (SCS-6) Fiber Reinforced Reaction-Formed Silicon Carbide Matrix Composites

Silicon carbide (SCS-6) fiber reinforced-reaction formed silicon carbide matrix composites were fabricated using NASA's reaction forming process. Silicon-2 at a percent of niobium alloy was used as an infiltrant instead of pure silicon to reduce the amount of free silicon in the matrix after reaction forming. The matrix primarily consists of silicon carbide with a bi-modal grain size distribution. Minority phases dispersed within the matrix are niobium disilicide (NbSi2), carbon and silicon. Fiber push-out tests on these composites determined a debond stress of approx. 67 MPa and a frictional stress of approx. 60 MPa. A typical four point flexural strength of the composite is 297 MPa (43.1 KSi). This composite shows tough behavior through fiber pull out.

Singh, Mrityunjay↗
NASA NTRS2019

Progress in Additively Manufactured Copper-Alloy GRCop-84, GRCop-42, and Bimetallic Combustion Chambers for Liquid Rocket Engines

Additive Manufacturing (AM) has significantly evolved over the last decade for use in the aerospace industry, particularly for liquid rocket engines. AM offers a considerable departure from traditional manufacturing to rapidly fabricate components with complex internal features. High performance liquid rocket engine combustion chambers that operate in a high heat flux environment are fabricated using a copper-alloy liner with a series of integral coolant channels. Copper-alloys provide the necessary conductivity and material strength for adequate design margins offering high performance without the need for film coolant. Copper-alloys present unique challenges to properly melt the powder in laser-based AM processes due to their high reflectivity and conductivity. Starting in 2014, NASA’s Marshall Space Flight Center (MSFC) and Glenn Research Center (GRC) have developed a process for AM of GRCop (Copper- Chrome-Niobium) alloys using Selective Laser Melting (SLM). GRCop, originally developed at GRC, is a high conductivity, high-strength, dispersion strengthened copper-alloy for use in high-temperature, high heat flux applications. NASA has completed significant material characterization and testing, along with hot-fire testing, to demonstrate that GRCop-42 and GRCop-84 alloys are suitable for use in combustion chambers. Additional development and testing has been completed on AM bimetallic chambers using GRCop-84 liners and superalloy jackets, fabricated using two Directed Energy Deposition (DED) processes: Electron Beam Freeform Fabrication (EBF³) and Blown Powder DED. NASA completed hot-fire testing on various AM chambers using GRCop-84, GRCop-42, and bimetallic chambers in Liquid Oxygen (LOX)/Hydrogen, LOX/Methane, and LOX/Kerosene propellants.

Gradl, Paul R.↗
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 NTRS1967

Protected, high-temperature connecting cable

Ceramic insulated, swaged stainless steel, sheathed, protective atmosphere cable admits electrical leads into an 1800 deg F air-environment test chamber. The cable has some bending capability and provides for nine niobium alloy conductors. An argon purge during the TIG weld closure protects internal wires from oxidation and embrittlement.

Engdahl, R. E.↗
NASA NTRS1971

Design and fabrication of brayton cycle solar heat receiver

A detail design and fabrication of a solar heat receiver using lithium fluoride as the heat storage material was completed. A gas flow analysis was performed to achieve uniform flow distribution within overall pressure drop limitations. Structural analyses and allowable design criteria were developed for anticipated environments such as launch, pressure containment, and thermal cycling. A complete heat receiver assembly was fabricated almost entirely from the refractory alloy, niobium-1% zirconium.

Mendelson, I.↗
NASA NTRS1989

Creep behavior of tungsten fiber reinforced niobium metal matrix composites

Tungsten fiber reinforced niobium metal matrix composites were evaluated for use in space nuclear power conversion systems. The composite panels were fabricated using the arc-spray monotape technique at the NASA Lewis Research Center. The creep behavior of W/Nb composite material was determined at 1400 and 1500 K in vacuum over a wide range of applied loads. The time to reach 1 percent strain, the time to rupture, and the minimum creep rate were measured. The W/Nb composites exceeded the properties of monolithic niobium alloys significantly even when compared on a strength to density basis. The effect of fiber orientation on the creep strength also was evaluated. Kirkendall void formation was observed at the fiber/matrix interface; the void distribution differed depending on the fiber orientation relative to the stress axis. A relationship was found between the fiber orientation and the creep strength.

Grobstein, T. L.↗
NASA NTRS1990

High-Temperature Creep Behavior Of Fiber-Reinforced Niobium

Study conducted to determine feasibility of using composite materials in advanced space power systems, described in 22-page report. Tungsten fibers reduce creep and mass in advanced power systems. Reinforcing niobium alloys with tungsten fibers increases their resistances to creep by factors of as much as 10.

Petrasek, Donald W.↗