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Melcher, Joh C. - One of the best experts on this subject based on the ideXlab platform.

  • Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
    2009
    Co-Authors: Melcher, Joh C., Allred, Jennife K.
    Abstract:

    Tests were conducted with the RS18 rocket engine using Liquid Oxygen (LO2) and Liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous Oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes

  • Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
    2009
    Co-Authors: Allred, Jennife K., Melcher, Joh C.
    Abstract:

    Tests were conducted with the RS-18 rocket engine using Liquid Oxygen (LO2) and Liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous Oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane

Allred, Jennife K. - One of the best experts on this subject based on the ideXlab platform.

  • Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
    2009
    Co-Authors: Melcher, Joh C., Allred, Jennife K.
    Abstract:

    Tests were conducted with the RS18 rocket engine using Liquid Oxygen (LO2) and Liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous Oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes

  • Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
    2009
    Co-Authors: Allred, Jennife K., Melcher, Joh C.
    Abstract:

    Tests were conducted with the RS-18 rocket engine using Liquid Oxygen (LO2) and Liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous Oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane

Gerhard Kroupa - One of the best experts on this subject based on the ideXlab platform.

  • laser ignition of a cryogenic thruster using a miniaturised nd yag laser
    Optics Express, 2013
    Co-Authors: Chiara Manfletti, Gerhard Kroupa
    Abstract:

    An experimental study has been conducted to assess the feasibility of implementing laser ignition in cryogenic reaction and control and orbital manouvering thrusters. A experimental thruster with a single-coaxial injector element combustion chamber for testing with Liquid Oxygen/gaseous hydrogen and Liquid Oxygen/gaseous methane was designed for this purpose. Mapping tests conducted using a standard table top laser revealed that the minimum incident energies required for 100% reliable laser plasma and laser ablation ignition of Liquid Oxygen/gaseous hydrogen are 72 mJ and 14.5 mJ respectively. In addition, the miniaturised HIPoLas® laser was mounted directly on the thruster and used as ignition system. This paper reports locations of energy deposition, levels of delivered energy and associated ignition probabilities obtained. The results indicate the feasibility of using a laser system for the direct ignition of reaction and control and orbital manouvering thrusters and highlight further investigations and developments necessary for the implementation of miniaturised laser systems for vacuum igntion of cryogenic propellants.

Chiara Manfletti - One of the best experts on this subject based on the ideXlab platform.

  • laser ignition of an experimental cryogenic reaction and control thruster ignition energies
    Journal of Propulsion and Power, 2014
    Co-Authors: Chiara Manfletti
    Abstract:

    An extensive experimental investigation has been conducted to assess the feasibility of implementing laser ignition in a cryogenic reaction and control system and orbital and maneuvering system thrusters. A 400 N experimental thruster, with a single-coaxial injector element, was designed for a steady-state pressure of 2 bar and was tested using two laser ignition systems: a tabletop laser as well as a miniaturized laser, which was mounted onto the thruster. Chamber pressures in the thruster before ignition varied from 200 mbar for Liquid Oxygen/gaseous methane to 450 mbar for Liquid Oxygen/gaseous hydrogen. The locations of energy deposition and the energy deposited were varied to determine the associated ignition probabilities for the propellant combinations Liquid Oxygen/gaseous hydrogen and Liquid Oxygen/gaseous methane. Plasma ignition and ablation ignition were both examined. For 100% ignition success, minimum energies measured were lower for Liquid Oxygen/gaseous hydrogen (72 mJ) than for Liquid oxy...

  • laser ignition of a cryogenic thruster using a miniaturised nd yag laser
    Optics Express, 2013
    Co-Authors: Chiara Manfletti, Gerhard Kroupa
    Abstract:

    An experimental study has been conducted to assess the feasibility of implementing laser ignition in cryogenic reaction and control and orbital manouvering thrusters. A experimental thruster with a single-coaxial injector element combustion chamber for testing with Liquid Oxygen/gaseous hydrogen and Liquid Oxygen/gaseous methane was designed for this purpose. Mapping tests conducted using a standard table top laser revealed that the minimum incident energies required for 100% reliable laser plasma and laser ablation ignition of Liquid Oxygen/gaseous hydrogen are 72 mJ and 14.5 mJ respectively. In addition, the miniaturised HIPoLas® laser was mounted directly on the thruster and used as ignition system. This paper reports locations of energy deposition, levels of delivered energy and associated ignition probabilities obtained. The results indicate the feasibility of using a laser system for the direct ignition of reaction and control and orbital manouvering thrusters and highlight further investigations and developments necessary for the implementation of miniaturised laser systems for vacuum igntion of cryogenic propellants.

Joel M. Stoltzfus - One of the best experts on this subject based on the ideXlab platform.

  • Liquid Oxygen rotating friction ignition testing of aluminum and titanium with monel and inconel for rocket engine propulsion system contamination investigation
    Journal of Astm International, 2009
    Co-Authors: Stephe Peralta, Keisa R Rosales, Joel M. Stoltzfus
    Abstract:

    Metallic contaminant was found in the Liquid Oxygen (LOX) pre-valve screen of the shuttle main engine propulsion system on two orbiter vehicles. To investigate the potential for an ignition, NASA Johnson Space Center White Sands Test Facility performed (modified) rotating friction ignition testing in LOX. This testing simulated a contaminant particle in the low-pressure Oxygen turbo pump (LPOTP) and the high-pressure Oxygen turbo pump (HPOTP) of the shuttle main propulsion system. Monel® K-500 and Inconel® 718 samples represented the LPOTP and HPOTP materials. Aluminum foil tape and titanium foil represented the contaminant particles. In both the Monel® and Inconel® material configurations, the aluminum foil tape samples did not ignite after 30 s of rubbing. In contrast, all of the titanium foil samples ignited regardless of the rubbing duration or material configuration. However, the titanium foil ignitions did not propagate to the Monel and Inconel materials. Based on these results, NASA determined that the risk of a fire occurring in the turbo pumps was acceptably low.

  • COMPOSITE MATERIAL COMPATIBILITY WITH Liquid Oxygen
    38th Structures Structural Dynamics and Materials Conference, 1997
    Co-Authors: Michael J Robinson, Mcdonnell Douglas Aerospace, Joel M. Stoltzfus, Thomas N. Owens
    Abstract:

    McDonnell Douglas Aerospace (MDA) and NASA are investigating the compatibility of composite materials and Liquid Oxygen (LO2). The objective is to determine if these advanced lightweight materials can be used to fabricate LO2 propellant tanks for launch vehicles. Rele- vant background on MDA's development of composite materials for Liquid hydrogen tanks is discussed. Mate- rial qualification standards for Liquid Oxygen service are reviewed. Hazards analyses on composite LO*, tank designs are discussed in which various ignition mecha- nisms are considered. Results are presented from a series of tests conducted to simulate potential ignition mechanisms in the presence of LO 2 and composite materials.