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Sample records for alloy-c-103

  1. Advanced Materials and Manufacturing for Low-Cost, High-Performance Liquid Rocket Combustion Chambers, Phase II Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Silicided niobium alloy (C103) combustion chambers have been used extensively in both NASA and DoD liquid rocket propulsion systems. Niobium alloys offer a good...

  2. Advanced Materials and Manufacturing for Low-Cost, High-Performance Liquid Rocket Combustion Chambers Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Silicided niobium alloy (C103) combustion chambers have been used extensively in both NASA and DoD liquid rocket propulsion systems. Niobium alloys offer a good...

  3. Effect of internal oxidation on the microstructure and mechanical properties of C-103 alloy

    Energy Technology Data Exchange (ETDEWEB)

    Sankar, M., E-mail: msankar_iitk@yahoo.co.in [Defence Metallurgical Research Laboratory, Hyderabad-500058, Andhra Pradesh (India); Baligidad, R.G.; Satyanarayana, D.V.V.; Gokhale, A.A. [Defence Metallurgical Research Laboratory, Hyderabad-500058, Andhra Pradesh (India)

    2013-07-01

    The effect of internal oxidation on the microstructure and mechanical properties of niobium alloy, C-103 has been investigated. Tensile specimens and test coupons of alloy containing different levels of oxygen (100–2500 ppm) were characterized with respect to microstructure and mechanical properties. It has been observed that for oxygen contents in the range ∼400–1000 ppm, hafnium oxide precipitated exclusively along the grain boundaries, while for oxygen content of ∼2500 ppm, precipitates formed both at the grain boundaries and within the grains near surface region of the alloy. The internal oxidation has resulted in embrittlement of the alloy resulting in considerable lowering of strength as well as ductility. Further, the strength and ductility are found to decrease progressively with the increase in average oxygen content of the alloy.

  4. Advanced materials for radiation-cooled rockets

    Science.gov (United States)

    Reed, Brian; Biaglow, James; Schneider, Steven

    1993-01-01

    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.

  5. Iridium-Coated Rhenium Radiation-Cooled Rockets

    Science.gov (United States)

    Reed, Brian D.; Biaglow, James A.; Schneider, Steven J.

    1997-01-01

    Radiation-cooled rockets are used for a range of low-thrust propulsion functions, including apogee insertion, attitude control, and repositioning of satellites, reaction control of launch vehicles, and primary propulsion for planetary space- craft. The key to high performance and long lifetimes for radiation-cooled rockets is the chamber temperature capability. The material system that is currently used for radiation-cooled rockets, a niobium alloy (C103) with a fused silica coating, has a maximum operating temperature of 1370 C. Temperature limitations of C103 rockets force the use of fuel film cooling, which degrades rocket performance and, in some cases, imposes a plume contamination issue from unburned fuel. A material system composed of a rhenium (Re) substrate and an iridium (Ir) coating has demonstrated operation at high temperatures (2200 C) and for long lifetimes (hours). The added thermal margin afforded by iridium-coated rhenium (Ir/Re) allows reduction or elimination of fuel film cooling. This, in turn, leads to higher performance and cleaner spacecraft environments. There are ongoing government- and industry-sponsored efforts to develop flight Ir/ Re engines, with the primary focus on 440-N, apogee insertion engines. Complementing these Ir/Re engine development efforts is a program to address specific concerns and fundamental characterization of the Ir/Re material system, including (1) development of Ir/Re rocket fabrication methods, (2) establishment of critical Re mechanical properly data, (3) development of reliable joining methods, and (4) characterization of Ir/Re life-limiting mechanisms.

  6. Attaching Thermocouples by Peening or Crimping

    Science.gov (United States)

    Murtland, Kevin; Cox, Robert; Immer, Christopher

    2006-01-01

    Two simple, effective techniques for attaching thermocouples to metal substrates have been devised for high-temperature applications in which attachment by such conventional means as welding, screws, epoxy, or tape would not be effective. The techniques have been used successfully to attach 0.005- in. (0.127-mm)-diameter type-S thermocouples to substrates of niobium alloy C-103 and stainless steel 416 for measuring temperatures up to 2,600 F (1,427 C). The techniques are equally applicable to other thermocouple and substrate materials. In the first technique, illustrated in the upper part of the figure, a hole slightly wider than twice the diameter of one thermocouple wire is drilled in the substrate. The thermocouple is placed in the hole, then the edge of the hole is peened in one or more places by use of a punch (see figure). The deformed material at the edge secures the thermocouple in the hole. In the second technique a hole is drilled as in the first technique, then an annular relief area is machined around the hole, resulting in structure reminiscent of a volcano in a crater. The thermocouple is placed in the hole as in the first technique, then the "volcano" material is either peened by use of a punch or crimped by use of sidecutters to secure the thermocouple in place. This second technique is preferable for very thin thermocouples [wire diameter .0.005 in. (.0.127 mm)] because standard peening poses a greater risk of clipping one or both of the thermocouple wires. These techniques offer the following advantages over prior thermocouple-attachment techniques: . Because these techniques involve drilling of very small holes, they are minimally invasive . an important advantage in that, to a first approximation, the thermal properties of surrounding areas are not appreciably affected. . These techniques do not involve introduction of any material, other than the substrate and thermocouple materials, that could cause contamination, could decompose, or oxidize