Science.gov

Sample records for gaseous rocket fuels

  1. Gaseous fuel nuclear reactor research

    NASA Technical Reports Server (NTRS)

    Schwenk, F. C.; Thom, K.

    1975-01-01

    Gaseous-fuel nuclear reactors are described; their distinguishing feature is the use of fissile fuels in a gaseous or plasma state, thereby breaking the barrier of temperature imposed by solid-fuel elements. This property creates a reactor heat source that may be able to heat the propellant of a rocket engine to 10,000 or 20,000 K. At this temperature level, gas-core reactors would provide the breakthrough in propulsion needed to open the entire solar system to manned and unmanned spacecraft. The possibility of fuel recycling makes possible efficiencies of up to 65% and nuclear safety at reduced cost, as well as high-thrust propulsion capabilities with specific impulse up to 5000 sec.

  2. Gaseous fuel nuclear reactor research

    NASA Technical Reports Server (NTRS)

    Schwenk, F. C.; Thom, K.

    1975-01-01

    Gaseous-fuel nuclear reactors are described; their distinguishing feature is the use of fissile fuels in a gaseous or plasma state, thereby breaking the barrier of temperature imposed by solid-fuel elements. This property creates a reactor heat source that may be able to heat the propellant of a rocket engine to 10,000 or 20,000 K. At this temperature level, gas-core reactors would provide the breakthrough in propulsion needed to open the entire solar system to manned and unmanned spacecraft. The possibility of fuel recycling makes possible efficiencies of up to 65% and nuclear safety at reduced cost, as well as high-thrust propulsion capabilities with specific impulse up to 5000 sec.

  3. Solid and Gaseous Fuels.

    ERIC Educational Resources Information Center

    Schultz, Hyman; And Others

    1989-01-01

    This review covers methods of sampling, analyzing, and testing coal, coke, and coal-derived solids and methods for the chemical, physical, and instrumental analyses of gaseous fuels. The review covers from October 1986, to September 1988. (MVL)

  4. Solid and Gaseous Fuels.

    ERIC Educational Resources Information Center

    Schultz, Hyman; And Others

    1989-01-01

    This review covers methods of sampling, analyzing, and testing coal, coke, and coal-derived solids and methods for the chemical, physical, and instrumental analyses of gaseous fuels. The review covers from October 1986, to September 1988. (MVL)

  5. Characteristics of response factors of coaxial gaseous rocket injectors

    NASA Technical Reports Server (NTRS)

    Janardan, B. A.; Daniel, B. R.; Zinn, B. T.

    1975-01-01

    The results of an experimental investigation undertaken to determine the frequency dependence of the response factors of various gaseous propellant rocket injectors subject to axial instabilities are presented. The injector response factors were determined, using the modified impedance-tube technique, under cold-flow conditions simulating those observed in unstable rocket motors. The tested injectors included a gaseous-fuel injector element, a gaseous-oxidizer injector element and a coaxial injector with both fuel and oxidizer elements. Emphasis was given to the determination of the dependence of the injector response factor upon the open-area ratio of the injector, the length of the injector orifice, and the pressure drop across the injector orifices. The measured data are shown to be in reasonable agreement with the corresponding injector response factor data predicted by the Feiler and Heidmann model.

  6. Gaseous fuel reactor research

    NASA Technical Reports Server (NTRS)

    Thom, K.; Schneider, R. T.

    1977-01-01

    The paper reviews studies dealing with the concept of a gaseous fuel reactor and describes the structure and plans of the current NASA research program of experiments on uranium hexafluoride systems and uranium plasma systems. Results of research into the basic properties of uranium plasmas and fissioning gases are reported. The nuclear pumped laser is described, and the main results of experiments with these devices are summarized.

  7. Improved hybrid rocket fuel

    NASA Technical Reports Server (NTRS)

    Dean, David L.

    1995-01-01

    McDonnell Douglas Aerospace, as part of its Independent R&D, has initiated development of a clean burning, high performance hybrid fuel for consideration as an alternative to the solid rocket thrust augmentation currently utilized by American space launch systems including Atlas, Delta, Pegasus, Space Shuttle, and Titan. It could also be used in single stage to orbit or as the only propulsion system in a new launch vehicle. Compared to solid propellants based on aluminum and ammonium perchlorate, this fuel is more environmentally benign in that it totally eliminates hydrogen chloride and aluminum oxide by products, producing only water, hydrogen, nitrogen, carbon oxides, and trace amounts of nitrogen oxides. Compared to other hybrid fuel formulations under development, this fuel is cheaper, denser, and faster burning. The specific impulse of this fuel is comparable to other hybrid fuels and is between that of solids and liquids. The fuel also requires less oxygen than similar hybrid fuels to produce maximum specific impulse, thus reducing oxygen delivery system requirements.

  8. Gaseous Fuel Injection Modeling using a Gaseous Sphere Injection Methodology

    SciTech Connect

    Hessel, R P; Aceves, S M; Flowers, D L

    2006-03-06

    The growing interest in gaseous fuels (hydrogen and natural gas) for internal combustion engines calls for the development of computer models for simulation of gaseous fuel injection, air entrainment and the ensuing combustion. This paper introduces a new method for modeling the injection and air entrainment processes for gaseous fuels. The model uses a gaseous sphere injection methodology, similar to liquid droplet in injection techniques used for liquid fuel injection. In this paper, the model concept is introduced and model results are compared with correctly- and under-expanded experimental data.

  9. Hydrogen and Gaseous Fuel Safety and Toxicity

    SciTech Connect

    Lee C. Cadwallader; J. Sephen Herring

    2007-06-01

    Non-traditional motor fuels are receiving increased attention and use. This paper examines the safety of three alternative gaseous fuels plus gasoline and the advantages and disadvantages of each. The gaseous fuels are hydrogen, methane (natural gas), and propane. Qualitatively, the overall risks of the four fuels should be close. Gasoline is the most toxic. For small leaks, hydrogen has the highest ignition probability and the gaseous fuels have the highest risk of a burning jet or cloud.

  10. The technique for Simulation of Transient Combustion Processes in the Rocket Engine Operating with Gaseous Fuel “Hydrogen and Oxygen”

    NASA Astrophysics Data System (ADS)

    Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

    2017-01-01

    The article describes the method for simulation of transient combustion processes in the rocket engine. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. Reactions mechanisms have been taken from several sources and verified. The method for converting ozone properties from the Shomate equation to the NASA-polynomial format was described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. Modeling difficulties with combustion model Finite Rate Chemistry, associated with a large scatter of reference data were identified and described. The way to generate the Flamelet library with CFX-RIF is described. Formulated adequate reaction mechanisms verified at a steady state have also been tested for transient simulation. The Flamelet combustion model was recognized as adequate for the transient mode. Integral parameters variation relates to the values obtained during stationary simulation. A cyclic irregularity of the temperature field, caused by precession of the vortex core, was detected in the chamber with the proposed simulation technique. Investigations of unsteady processes of rocket engines including the processes of ignition were proposed as the area for application of the described simulation technique.

  11. Experimental Evaluation of a Subscale Gaseous Hydrogen/Gaseous Oxygen Coaxial Rocket Injector

    NASA Astrophysics Data System (ADS)

    Smith, Timothy D.; Klem, Mark D.; Breisacher, Kevin J.; Farhangi, Shahram; Sutton, Robert

    2002-11-01

    The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested.

  12. Experimental Evaluation of a Subscale Gaseous Hydrogen/gaseous Oxygen Coaxial Rocket Injector

    NASA Technical Reports Server (NTRS)

    Smith, Timothy D.; Klem, Mark D.; Breisacher, Kevin J.; Farhangi, Shahram; Sutton, Robert

    2002-01-01

    The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested.

  13. Performance Capability of Single-Cavity Vortex Gaseous Nuclear Rockets

    NASA Technical Reports Server (NTRS)

    Ragsdale, Robert G.

    1963-01-01

    An analysis was made to determine the maximum powerplant thrust-to-weight ratio possible with a single-cavity vortex gaseous reactor in which all the hydrogen propellant must diffuse through a fuel-rich region. An assumed radial temperature profile was used to represent conduction, convection, and radiation heat-transfer effects. The effect of hydrogen property changes due to dissociation and ionization was taken into account in a hydrodynamic computer program. It is shown that, even for extremely optimistic assumptions of reactor criticality and operating conditions, such a system is limited to reactor thrust-to-weight ratios of about 1.2 x 10(exp -3) for laminar flow. For turbulent flow, the maximum thrust-to-weight ratio is less than 10(exp -3). These low thrusts result from the fact that the hydrogen flow rate is limited by the diffusion process. The performance of a gas-core system with a specific impulse of 3000 seconds and a powerplant thrust-to-weight ratio of 10(exp -2) is shown to be equivalent to that of a 1000-second advanced solid-core system. It is therefore concluded that a single-cavity vortex gaseous reactor in which all the hydrogen must diffuse through the nuclear fuel is a low-thrust device and offers no improvement over a solid-core nuclear-rocket engine. To achieve higher thrust, additional hydrogen flow must be introduced in such a manner that it will by-pass the nuclear fuel. Obviously, such flow must be heated by thermal radiation. An illustrative model of a single-cavity vortex system employing supplementary flow of hydrogen through the core region is briefly examined. Such a system appears capable of thrust-to-weight ratios of approximately 1 to 10. For a high-impulse engine, this capability would be a considerable improvement over solid-core performance. Limits imposed by thermal radiation heat transfer to cavity walls are acknowledged but not evaluated. Alternate vortex concepts that employ many parallel vortices to achieve higher

  14. Design of a Novel Gaseous Hydrogen-Oxygen Rocket Injector Element

    NASA Technical Reports Server (NTRS)

    Glenn, Dennis

    1999-01-01

    An overview of activities supporting the design of a gaseous hydrogen-oxygen rocket injector element is presented in viewgraph form. The purpose of the research was to find a viable design for a rocket gas-gas injector that mixes fuel and oxidizer thoroughly and quickly. Computational fluid dynamics analyses were used with reacting flow to evaluate design options for mixing, temperature distribution, and combustion efficiency. A design was found that is an improvement over designs derived from liquid systems and is far better than traditional shear-coax.

  15. Investigation of gaseous nuclear rocket technology

    NASA Technical Reports Server (NTRS)

    Kendall, J. S.

    1972-01-01

    The experimental and theoretical investigations conducted during the period from September 1969 through September 1972 are reported which were directed toward obtaining information necessary to determine the feasibility of the full-scale nuclear light bulb engine, and of small-scale nuclear tests involving fissioning uranium plasmas in a unit cell installed in a driver reactor, such as the Nuclear Furnace. Emphasis was placed on development of RF simulations of conditions expected in nuclear tests in the Nuclear Furnace. The work included investigations of the following: (1) the fluid mechanics and containment characteristics of one-component and two-component vortex flows, both unheated and RF-induction heated; (2) heating of particle-seeded streams by thermal radiation from a dc arc to simulate propellant heating; (3) condensation and separation phenomena for metal-vapor/heated-gas mixtures to provide information for conceptual designs of components of fuel exhaust and recycle systems; (4) the characteristics of the radiant energy spectrum emitted from the fuel region, with emphasis on definition of fuel and buffer-gas region seed systems to reduce the ultraviolet radiation emitted from the nuclear fuel; and (5) the effects of nuclear radiation on the optical transmission characteristics of transparent materials.

  16. Gaseous fuel reactors for power systems

    NASA Technical Reports Server (NTRS)

    Kendall, J. S.; Rodgers, R. J.

    1977-01-01

    Gaseous-fuel nuclear reactors have significant advantages as energy sources for closed-cycle power systems. The advantages arise from the removal of temperature limits associated with conventional reactor fuel elements, the wide variety of methods of extracting energy from fissioning gases, and inherent low fissile and fission product in-core inventory due to continuous fuel reprocessing. Example power cycles and their general performance characteristics are discussed. Efficiencies of gaseous fuel reactor systems are shown to be high with resulting minimal environmental effects. A technical overview of the NASA-funded research program in gaseous fuel reactors is described and results of recent tests of uranium hexafluoride (UF6)-fueled critical assemblies are presented.

  17. Heterogeneous fuel for hybrid rocket

    NASA Technical Reports Server (NTRS)

    Stickler, David B. (Inventor)

    1996-01-01

    Heterogeneous fuel compositions suitable for use in hybrid rocket engines and solid-fuel ramjet engines, The compositions include mixtures of a continuous phase, which forms a solid matrix, and a dispersed phase permanently distributed therein. The dispersed phase or the matrix vaporizes (or melts) and disperses into the gas flow much more rapidly than the other, creating depressions, voids and bumps within and on the surface of the remaining bulk material that continuously roughen its surface, This effect substantially enhances heat transfer from the combusting gas flow to the fuel surface, producing a correspondingly high burning rate, The dispersed phase may include solid particles, entrained liquid droplets, or gas-phase voids having dimensions roughly similar to the displacement scale height of the gas-flow boundary layer generated during combustion.

  18. Study of Rapid-Regression Liquefying Hybrid Rocket Fuels

    NASA Technical Reports Server (NTRS)

    Zilliac, Greg; DeZilwa, Shane; Karabeyoglu, M. Arif; Cantwell, Brian J.; Castellucci, Paul

    2004-01-01

    A report describes experiments directed toward the development of paraffin-based hybrid rocket fuels that burn at regression rates greater than those of conventional hybrid rocket fuels like hydroxyl-terminated butadiene. The basic approach followed in this development is to use materials such that a hydrodynamically unstable liquid layer forms on the melting surface of a burning fuel body. Entrainment of droplets from the liquid/gas interface can substantially increase the rate of fuel mass transfer, leading to surface regression faster than can be achieved using conventional fuels. The higher regression rate eliminates the need for the complex multi-port grain structures of conventional solid rocket fuels, making it possible to obtain acceptable performance from single-port structures. The high-regression-rate fuels contain no toxic or otherwise hazardous components and can be shipped commercially as non-hazardous commodities. Among the experiments performed on these fuels were scale-up tests using gaseous oxygen. The data from these tests were found to agree with data from small-scale, low-pressure and low-mass-flux laboratory tests and to confirm the expectation that these fuels would burn at high regression rates, chamber pressures, and mass fluxes representative of full-scale rocket motors.

  19. THE LIQUID AND GASEOUS FUEL DISTRIBUTION SYSTEM

    EPA Science Inventory

    The report describes the national liquid and gaseous fuel distribution system. he study leading to the report was performed as part of an effort to better understand emissions of volatile organic compounds from the fuel distribution system. he primary, secondary, and tertiary seg...

  20. THE LIQUID AND GASEOUS FUEL DISTRIBUTION SYSTEM

    EPA Science Inventory

    The report describes the national liquid and gaseous fuel distribution system. he study leading to the report was performed as part of an effort to better understand emissions of volatile organic compounds from the fuel distribution system. he primary, secondary, and tertiary seg...

  1. Development of high performance hybrid rocket fuels

    NASA Astrophysics Data System (ADS)

    Zaseck, Christopher R.

    . In order to examine paraffin/additive combustion in a motor environment, I conducted experiments on well characterized aluminum based additives. In particular, I investigate the influence of aluminum, unpassivated aluminum, milled aluminum/polytetrafluoroethylene (PTFE), and aluminum hydride on the performance of paraffin fuels for hybrid rocket propulsion. I use an optically accessible combustor to examine the performance of the fuel mixtures in terms of characteristic velocity efficiency and regression rate. Each combustor test consumes a 12.7 cm long, 1.9 cm diameter fuel strand under 160 kg/m 2s of oxygen at up to 1.4 MPa. The experimental results indicate that the addition of 5 wt.% 30 mum or 80 nm aluminum to paraffin increases the regression rate by approximately 15% compared to neat paraffin grains. At higher aluminum concentrations and nano-scale particles sizes, the increased melt layer viscosity causes slower regression. Alane and Al/PTFE at 12.5 wt.% increase the regression of paraffin by 21% and 32% respectively. Finally, an aging study indicates that paraffin can protect air and moisture sensitive particles from oxidation. The opposed burner and aluminum/paraffin hybrid rocket experiments show that additives can alter bulk fuel properties, such as viscosity, that regulate entrainment. The general effect of melt layer properties on the entrainment and regression rate of paraffin is not well understood. Improved understanding of how solid additives affect the properties and regression of paraffin is essential to maximize performance. In this document I investigate the effect of melt layer properties on paraffin regression using inert additives. Tests are performed in the optical cylindrical combustor at ˜1 MPa under a gaseous oxygen mass flux of ˜160 kg/m2s. The experiments indicate that the regression rate is proportional to mu0.08rho 0.38kappa0.82. In addition, I explore how to predict fuel viscosity, thermal conductivity, and density prior to testing

  2. Gaseous-fuel safety assessment. Status report

    SciTech Connect

    Krupka, M.C.; Edeskuty, F.J.; Bartlit, J.R.; Williamson, K.D. Jr.

    1982-01-01

    The Los Alamos National Laboratory, in support of studies sponsored by the Office of Vehicle and Engine Research and Development in the US Department of Energy, has undertaken a safety assessment of selected gaseous fuels for use in light automotive transportation. The purpose is to put into perspective the hazards of these fuels relative to present day fuels and delineated criteria for their safe handling. Fuels include compressed and liquified natural gas (CNG and LNG), liquefied petroleum gas (LPG), and for reference gasoline and diesel. This paper is a program status report. To date, physicochemical property data and general petroleum and transportation information were compiled; basic hazards defined; alternative fuels were safety-ranked based on technical properties alone; safety data and vehicle accident statistics reviewed; and accident scenarios selected for further analysis. Methodology for such analysis is presently under consideration.

  3. Heat transfer analysis of fuel assemblies in a heterogeneous gas core nuclear rocket

    NASA Technical Reports Server (NTRS)

    Watanabe, Yoichi; Appelbaum, Jacob; Diaz, Nils; Maya, Isaac

    1991-01-01

    Heat transfer problems of a heterogeneous gaseous core nuclear rocket were studied. The reactor core consists of 1.5-m long hexagonal fuel assemblies filled with pressurized uranium tetrafluoride (UF4) gas. The fuel gas temperature ranges from 3500 to 7000 K at a nominal operating condition of 40 atm. Each fuel assembly has seven coolant tubes, through which hydrogen propellant flows. The propellant temperature is not constrained by the fuel temperature but by the maximum temperature of the graphite coolant tube. For a core achieving a fission power density of 1000 MW/cu m, the propellant core exit temperature can be as high as 3200 K. The physical size of a 1250 MW gaseous core nuclear rocket is comparable with that of a NERVA-type solid core nuclear rocket. The engine can deliver a specific impulse of 1020 seconds and a thrust of 330 kN.

  4. Heat transfer analysis of fuel assemblies in a heterogeneous gas core nuclear rocket

    NASA Technical Reports Server (NTRS)

    Watanabe, Yoichi; Appelbaum, Jacob; Diaz, Nils; Maya, Isaac

    1991-01-01

    Heat transfer problems of a heterogeneous gaseous core nuclear rocket were studied. The reactor core consists of 1.5-m long hexagonal fuel assemblies filled with pressurized uranium tetrafluoride (UF4) gas. The fuel gas temperature ranges from 3500 to 7000 K at a nominal operating condition of 40 atm. Each fuel assembly has seven coolant tubes, through which hydrogen propellant flows. The propellant temperature is not constrained by the fuel temperature but by the maximum temperature of the graphite coolant tube. For a core achieving a fission power density of 1000 MW/cu m, the propellant core exit temperature can be as high as 3200 K. The physical size of a 1250 MW gaseous core nuclear rocket is comparable with that of a NERVA-type solid core nuclear rocket. The engine can deliver a specific impulse of 1020 seconds and a thrust of 330 kN.

  5. Combustion characteristics of alternative gaseous fuels

    SciTech Connect

    Park, O.; Veloo, Peter S.; Liu, N.; Egolfopoulos, Fokion N.

    2011-01-01

    Fundamental flame properties of mixtures of air with hydrogen, carbon monoxide, and C{sub 1}–C{sub 4} saturated hydrocarbons were studied both experimentally and numerically. The fuel mixtures were chosen in order to simulate alternative gaseous fuels and to gain insight into potential kinetic couplings during the oxidation of fuel mixtures. The studies included the use of the counterflow configuration for the determination of laminar flame speeds, as well as extinction and ignition limits of premixed flames. The experiments were modeled using the USC Mech II kinetic model. It was determined that when hydrocarbons are added to hydrogen flames as additives, flame ignition, propagation, and extinction are affected in a counterintuitive manner. More specifically, it was found that by substituting methane by propane or n-butane in hydrogen flames, the reactivity of the mixture is reduced both under pre-ignition and vigorous burning conditions. This behavior stems from the fact that propane and n-butane produce higher amounts of methyl radicals that can readily recombine with atomic hydrogen and reduce thus the rate of the H + O{sub 2} → O + OH branching reaction. The kinetic model predicts closely the experimental data for flame propagation and extinction for various fuel mixtures and pressures, and for various amounts of carbon dioxide in the fuel blend. On the other hand, it underpredicts, in general, the ignition temperatures.

  6. Deposit formation in hydrocarbon rocket fuels

    NASA Technical Reports Server (NTRS)

    Roback, R.; Szetela, E. J.; Spadaccini, L. J.

    1981-01-01

    An experimental program was conducted to study deposit formation in hydrocarbon fuels under flow conditions that exist in high-pressure, rocket engine cooling systems. A high pressure fuel coking test apparatus was designed and developed and was used to evaluate thermal decomposition (coking) limits and carbon deposition rates in heated copper tubes for two hydrocarbon rocket fuels, RP-1 and commercial-grade propane. Tests were also conducted using JP-7 and chemically-pure propane as being representative of more refined cuts of the baseline fuels. A parametric evaluation of fuel thermal stability was performed at pressures of 136 atm to 340 atm, bulk fuel velocities in the range 6 to 30 m/sec, and tube wall temperatures in the range 422 to 811 K. Results indicated that substantial deposit formation occurs with RP-1 fuel at wall temperatures between 600 and 800 K, with peak deposit formation occurring near 700 K. No improvements were obtained when deoxygenated JP-7 fuel was substituted for RP-1. The carbon deposition rates for the propane fuels were generally higher than those obtained for either of the kerosene fuels at any given wall temperature. There appeared to be little difference between commercial-grade and chemically-pure propane with regard to type and quantity of deposit. Results of tests conducted with RP-1 indicated that the rate of deposit formation increased slightly with pressure over the range 136 atm to 340 atm. Finally, lating the inside wall of the tubes with nickel was found to significantly reduce carbon deposition rates for RP-1 fuel.

  7. Nuclear rocket using indigenous Martian fuel NIMF

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert

    1991-01-01

    In the 1960's, Nuclear Thermal Rocket (NTR) engines were developed and ground tested capable of yielding isp of up to 900 s at thrusts up to 250 klb. Numerous trade studies have shown that such traditional hydrogen fueled NTR engines can reduce the inertial mass low earth orbit (IMLEO) of lunar missions by 35 percent and Mars missions by 50 to 65 percent. The same personnel and facilities used to revive the hydrogen NTR can also be used to develop NTR engines capable of using indigenous Martian volatiles as propellant. By putting this capacity of the NTR to work in a Mars descent/acent vehicle, the Nuclear rocket using Indigenous Martian Fuel (NIMF) can greatly reduce the IMLEO of a manned Mars mission, while giving the mission unlimited planetwide mobility.

  8. Liquid fuel injection elements for rocket engines

    NASA Technical Reports Server (NTRS)

    Cox, George B., Jr. (Inventor)

    1993-01-01

    Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement. An injection element for a liquid propellant rocket engine of the bipropellant type is provided which includes tangential fuel metering orifices, and a plurality of oxidizer tube injection elements whose injection tubes are also provided with tangential oxidizer entry slots and internal reed valves.

  9. Liquid fuel injection elements for rocket engines

    NASA Astrophysics Data System (ADS)

    Cox, George B., Jr.

    1993-01-01

    Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement. An injection element for a liquid propellant rocket engine of the bipropellant type is provided herein which includes tangential fuel metering orifices, and a plurality of oxidizer tube injection elements whose injection tubes are also provided with tangential oxidizer entry slots and internal reed valves.

  10. Liquid fuel injection elements for rocket engines

    NASA Astrophysics Data System (ADS)

    Cox, George B., Jr.

    1993-11-01

    Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement. An injection element for a liquid propellant rocket engine of the bipropellant type is provided which includes tangential fuel metering orifices, and a plurality of oxidizer tube injection elements whose injection tubes are also provided with tangential oxidizer entry slots and internal reed valves.

  11. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 49 Transportation 6 2012-10-01 2012-10-01 false Gallon Equivalents for Gaseous Fuels. 538.8 Section 538.8 Transportation Other Regulations Relating to Transportation (Continued) NATIONAL HIGHWAY TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL...

  12. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... 49 Transportation 6 2013-10-01 2013-10-01 false Gallon Equivalents for Gaseous Fuels. 538.8 Section 538.8 Transportation Other Regulations Relating to Transportation (Continued) NATIONAL HIGHWAY TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL...

  13. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 49 Transportation 6 2014-10-01 2014-10-01 false Gallon Equivalents for Gaseous Fuels. 538.8 Section 538.8 Transportation Other Regulations Relating to Transportation (Continued) NATIONAL HIGHWAY TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL...

  14. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 49 Transportation 6 2011-10-01 2011-10-01 false Gallon Equivalents for Gaseous Fuels. 538.8 Section 538.8 Transportation Other Regulations Relating to Transportation (Continued) NATIONAL HIGHWAY TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL...

  15. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... 49 Transportation 6 2010-10-01 2010-10-01 false Gallon Equivalents for Gaseous Fuels. 538.8 Section 538.8 Transportation Other Regulations Relating to Transportation (Continued) NATIONAL HIGHWAY TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL...

  16. Computational simulation of liquid fuel rocket injectors

    NASA Technical Reports Server (NTRS)

    Landrum, D. Brian

    1994-01-01

    A major component of any liquid propellant rocket is the propellant injection system. Issues of interest include the degree of liquid vaporization and its impact on the combustion process, the pressure and temperature fields in the combustion chamber, and the cooling of the injector face and chamber walls. The Finite Difference Navier-Stokes (FDNS) code is a primary computational tool used in the MSFC Computational Fluid Dynamics Branch. The branch has dedicated a significant amount of resources to development of this code for prediction of both liquid and solid fuel rocket performance. The FDNS code is currently being upgraded to include the capability to model liquid/gas multi-phase flows for fuel injection simulation. An important aspect of this effort is benchmarking the code capabilities to predict existing experimental injection data. The objective of this MSFC/ASEE Summer Faculty Fellowship term was to evaluate the capabilities of the modified FDNS code to predict flow fields with liquid injection. Comparisons were made between code predictions and existing experimental data. A significant portion of the effort included a search for appropriate validation data. Also, code simulation deficiencies were identified.

  17. Hydrochloric acid aerosol and gaseous hydrogen chloride partitioning in a cloud contaminated by solid rocket exhaust

    NASA Technical Reports Server (NTRS)

    Sebacher, D. I.; Bendura, R. J.; Wornom, D. E.

    1980-01-01

    Partitioning of hydrogen chloride between hydrochloric acid aerosol and gaseous HCl in the lower atmosphere was experimentally investigated in a solid rocket exhaust cloud diluted with humid ambient air. Airborne measurements were obtained of gaseous HCl, total HCl, relative humidity and temperature to evaluate the conditions under which aerosol formation occurs in the troposphere in the presence of hygroscopic HCl vapor. Equilibrium predictions of HCl aerosol formation accurately predict the measured HCl partitioning over a range of total HCl concentrations from 0.6 to 16 ppm.

  18. Hydrochloric acid aerosol and gaseous hydrogen chloride partitioning in a cloud contaminated by solid rocket exhaust

    NASA Technical Reports Server (NTRS)

    Sebacher, D. I.; Bendura, R. J.; Wornom, D. E.

    1980-01-01

    Partitioning of hydrogen chloride between hydrochloric acid aerosol and gaseous HCl in the lower atmosphere was experimentally investigated in a solid rocket exhaust cloud diluted with humid ambient air. Airborne measurements were obtained of gaseous HCl, total HCl, relative humidity and temperature to evaluate the conditions under which aerosol formation occurs in the troposphere in the presence of hygroscopic HCl vapor. Equilibrium predictions of HCl aerosol formation accurately predict the measured HCl partitioning over a range of total HCl concentrations from 0.6 to 16 ppm.

  19. Injector for liquid fueled rocket engine

    NASA Technical Reports Server (NTRS)

    Cornelius, Charles S. (Inventor); Myers, W. Neill (Inventor); Shadoan, Michael David (Inventor); Sparks, David L. (Inventor)

    2000-01-01

    An injector for liquid fueled rocket engines wherein a generally flat core having a frustoconical dome attached to one side of the core to serve as a manifold for a first liquid, with the core having a generally circular configuration having an axis. The other side of the core has a plurality of concentric annular first slots and a plurality of annular concentric second slots alternating with the first slots, the second slots having a greater depth than said first slots. A bore extends through the core for inletting a second liquid into said core, the bore intersecting the second slots to feed the second liquid into the second slots. The core also has a plurality of first passageways leading from the manifold to the first annular slots for feeding the first liquid into said first slots. A faceplate brazed to said other side of the core is provided with apertures extending from the first and second slots through said face plate, these apertures being positioned to direct fuel and liquid oxygen into contact with each other in the combustion chamber. The first liquid may be liquid oxygen and the second liquid may be kerosene or liquid hydrogen.

  20. Design handbook for gaseous fuel engine injectors and combustion chambers

    NASA Technical Reports Server (NTRS)

    Calhoon, D. F.; Ito, I.; Kors, D. L.

    1973-01-01

    Results of investigation of injection, mixing, and combustion processes using gaseous fuels and oxidizers have been summarized in handbook presenting succinct design procedures for injectors and methods for estimating combustion efficiency, chamber heat flux and stability characteristics. Handbook presents two approaches to injector and combustion chamber design: empirical and analytical.

  1. Grooved Fuel Rings for Nuclear Thermal Rocket Engines

    NASA Technical Reports Server (NTRS)

    Emrich, William

    2009-01-01

    An alternative design concept for nuclear thermal rocket engines for interplanetary spacecraft calls for the use of grooved-ring fuel elements. Beyond spacecraft rocket engines, this concept also has potential for the design of terrestrial and spacecraft nuclear electric-power plants. The grooved ring fuel design attempts to retain the best features of the particle bed fuel element while eliminating most of its design deficiencies. In the grooved ring design, the hydrogen propellant enters the fuel element in a manner similar to that of the Particle Bed Reactor (PBR) fuel element.

  2. Low NOx heavy fuel combustor concept program addendum: Low/mid heating value gaseous fuel evaluation

    NASA Technical Reports Server (NTRS)

    Novick, A. S.; Troth, D. L.

    1982-01-01

    The combustion performance of a rich/quench/lean (RQL) combustor was evaluated when operated on low and mid heating value gaseous fuels. Two synthesized fuels were prepared having lower heating values of 10.2 MJ/cu m. (274 Btu/scf) and 6.6 MJ/cu m (176 Btu/scf). These fuels were configured to be representative of actual fuels, being composed primarily of nitrogen, hydrogen, carbon monoxide, and carbon dioxide. A liquid fuel air assist fuel nozzle was modified to inject both of the gaseous fuels. The RQL combustor liner was not changed from the configuration used when the liquid fuels were tested. Both gaseous fuels were tested over a range of power levels from 50 percent load to maximum rated power of the DDN Model 570-K industrial gas turbine engine. Exhaust emissions were recorded for four power level at several rich zone equivalence ratios to determine NOx sensitivity to the rich zone operating point. For the mid Btu heating value gas, ammonia was added to the fuel to simulate a fuel bound nitrogen type gaseous fuel. Results at the testing showed that for the low heating value fuel NOx emissions were all below 20 ppmc and smoke was below a 10 smoke number. For the mid heating value fuel, NOx emissions were in the 50 to 70 ppmc range with the smoke below a 10 smoke number.

  3. SpaceX rocket fuel plan under scrutiny

    NASA Astrophysics Data System (ADS)

    Gwynne, Peter

    2016-12-01

    NASA's International Space Station advisory committee has raised concerns about SpaceX's plans to fuel rockets that are used to ferry astronauts to the International Space Station (ISS) while the crew is onboard.

  4. Combustion characteristics of hydrogen-carbon monoxide based gaseous fuels

    NASA Technical Reports Server (NTRS)

    White, D. J.; Kubasco, A. J.; Lecren, R. T.; Notardonato, J. J.

    1982-01-01

    The results of trials with a staged combustor designed to use coal-derived gaseous fuels and reduce the NO(x) emissions from nitrogen-bound fuels to 75 ppm and 37 ppm without bound nitrogen in 15% O2 are reported. The combustor was outfitted with primary zone regenerative cooling, wherein the air cooling the primary zone was passed into the combustor at 900 F and mixed with the fuel. The increase in the primary air inlet temperature eliminated flashback and autoignition, lowered the levels of CO, unburned hydrocarbons, and smoke, and kept combustion efficiencies to the 99% level. The combustor was also equipped with dual fuel injection to test various combinations of liquid/gas fuel mixtures. Low NO(x) emissions were produced burning both Lurgi and Winkler gases, regardless of the inlet pressure and temperature conditions. Evaluation of methanation of medium energy gases is recommended for providing a fuel with low NO(x) characteristics.

  5. Combustion characteristics of hydrogen-carbon monoxide based gaseous fuels

    NASA Technical Reports Server (NTRS)

    White, D. J.; Kubasco, A. J.; Lecren, R. T.; Notardonato, J. J.

    1982-01-01

    The results of trials with a staged combustor designed to use coal-derived gaseous fuels and reduce the NO(x) emissions from nitrogen-bound fuels to 75 ppm and 37 ppm without bound nitrogen in 15% O2 are reported. The combustor was outfitted with primary zone regenerative cooling, wherein the air cooling the primary zone was passed into the combustor at 900 F and mixed with the fuel. The increase in the primary air inlet temperature eliminated flashback and autoignition, lowered the levels of CO, unburned hydrocarbons, and smoke, and kept combustion efficiencies to the 99% level. The combustor was also equipped with dual fuel injection to test various combinations of liquid/gas fuel mixtures. Low NO(x) emissions were produced burning both Lurgi and Winkler gases, regardless of the inlet pressure and temperature conditions. Evaluation of methanation of medium energy gases is recommended for providing a fuel with low NO(x) characteristics.

  6. Very Low Thrust Gaseous Oxygen-hydrogen Rocket Engine Ignition Technology

    NASA Technical Reports Server (NTRS)

    Bjorklund, Roy A.

    1983-01-01

    An experimental program was performed to determine the minimum energy per spark for reliable and repeatable ignition of gaseous oxygen (GO2) and gaseous hydrogen (GH2) in very low thrust 0.44 to 2.22-N (0.10 to 0.50-lb sub f) rocket engines or spacecraft and satellite attitude control systems (ACS) application. Initially, the testing was conducted at ambient conditions, with the results subsequently verified under vacuum conditions. An experimental breadboard electrical exciter that delivered 0.2 to 0.3 mj per spark was developed and demonstrated by repeated ignitions of a 2.22-N (0.50-lb sub f) thruster in a vacuum chamber with test durations up to 30 min.

  7. The use of gaseous fuels mixtures for SI engines propulsion

    NASA Astrophysics Data System (ADS)

    Flekiewicz, M.; Kubica, G.

    2016-09-01

    Paper presents results of SI engine tests, carried on for different gaseous fuels. Carried out analysis made it possible to define correlation between fuel composition and engine operating parameters. Tests covered various gaseous mixtures: of methane and hydrogen and LPG with DME featuring different shares. The first group, considered as low carbon content fuels can be characterized by low CO2 emissions. Flammability of hydrogen added in those mixtures realizes the function of combustion process activator. That is why hydrogen addition improves the energy conversion by about 3%. The second group of fuels is constituted by LPG and DME mixtures. DME mixes perfectly with LPG, and differently than in case of other hydrocarbon fuels consists also of oxygen makes the stoichiometric mixture less oxygen demanding. In case of this fuel an improvement in engine volumetric and overall engine efficiency has been noticed, when compared to LPG. For the 11% DME share in the mixture an improvement of 2% in the efficiency has been noticed. During the tests standard CNG/LPG feeding systems have been used, what underlines utility value of the research. The stand tests results have been followed by combustion process simulation including exhaust forming and charge exchange.

  8. Non-homogeneous hybrid rocket fuel for enhanced regression rates utilizing partial entrainment

    NASA Astrophysics Data System (ADS)

    Boronowsky, Kenny

    A concept was developed and tested to enhance the performance and regression rate of hydroxyl terminated polybutadiene (HTPB), a commonly used hybrid rocket fuel. By adding small nodules of paraffin into the HTPB fuel, a non-homogeneous mixture was created resulting in increased regression rates. The goal was to develop a fuel with a simplified single core geometry and a tailorable regression rate. The new fuel would benefit from the structural stability of HTPB yet not suffer from the large void fraction representative of typical HTPB core geometries. Regression rates were compared between traditional HTPB single core grains, 85% HTPB mixed with 15% (by weight) paraffin cores, 70% HTPB mixed with 30% paraffin cores, and plain paraffin single core grains. Each fuel combination was tested at oxidizer flow rates, ranging from 0.9 - 3.3 g/s of gaseous oxygen, in a small scale hybrid test rocket and average regression rates were measured. While large uncertainties were present in the experimental setup, the overall data showed that the regression rate was enhanced as paraffin concentration increased. While further testing would be required at larger scales of interest, the trends are encouraging. Inclusion of paraffin nodules in the HTPB grain may produce a greater advantage than other more noxious additives in current use. In addition, it may lead to safer rocket motors with higher integrated thrust due to the decreased void fraction.

  9. Water rocket - Electrolysis propulsion and fuel cell power

    SciTech Connect

    Carter, P H; Dittman, M D; Kare, J T; Militsky, F; Myers, B; Weisberg, A H

    1999-07-24

    Water Rocket is the collective name for an integrated set of technologies that offer new options for spacecraft propulsion, power, energy storage, and structure. Low pressure water stored on the spacecraft is electrolyzed to generate, separate, and pressurize gaseous hydrogen and oxygen. These gases, stored in lightweight pressure tanks, can be burned to generate thrust or recombined to produce electric power. As a rocket propulsion system, Water Rocket provides the highest feasible chemical specific impulse (-400 seconds). Even higher specific impulse propulsion can be achieved by combining Water Rocket with other advanced propulsion technologies, such as arcjet or electric thrusters. With innovative pressure tank technology, Water Rocket's specific energy [Wh/kg] can exceed that of the best foreseeable batteries by an order of magnitude, and the tanks can often serve as vehicle structural elements. For pulsed power applications, Water Rocket propellants can be used to drive very high power density generators, such as MHD devices or detonation-driven pulse generators. A space vehicle using Water Rocket propulsion can be totally inert and non-hazardous during assembly and launch. These features are particularly important for the timely development and flight qualification of new classes of spacecraft, such as microsats, nanosats, and refuelable spacecraft.

  10. NOx formation in combustion of gaseous fuel in ejection burner

    NASA Astrophysics Data System (ADS)

    Rimár, Miroslav; Kulikov, Andrii

    2016-06-01

    The aim of this work is to prepare model for researching of the formation in combustion of gaseous fuels. NOx formation is one of the main ecological problems nowadays as nitrogen oxides is one of main reasons of acid rains. The ANSYS model was designed according to the calculation to provide full combustion and good mixing of the fuel and air. The current model is appropriate to research NOx formation and the influence of the different principles of NOx reduction method. Applying of designed model should spare both time of calculations and research and also money as you do not need to measure the burner characteristics.

  11. Combustion of solid fuel slabs with gaseous oxygen in a hybrid motor analog

    NASA Technical Reports Server (NTRS)

    Chiaverini, Martin J.; Harting, George C.; Lu, Yeu-Cherng; Kuo, Kenneth K.; Serin, Nadir; Johnson, David K.

    1995-01-01

    Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated Polybutadiene) fuel cross-linked with diisocyanate was burned with gaseous oxygen (GOX) under various operating conditions. Large-amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed-line system and combustion chamber, the pressure oscillations were drastically reduced from plus or minus 20% of the localized mean pressure to an acceptable range of plus or minus 1.5%. Embedded fine--wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading edge region, the subsurface thermal wave profiles in the upstream locations are thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real-time X-ray radiography and ultrasonic pulse echo techniques were used to determine the instantaneous web thicknesses and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented. Several tests were conducted using, simultaneously, one translucent fuel slab and one fuel slab processed with carbon black powder. The addition of carbon black did not affect the measured regression rates or

  12. Experimental and Detailed Numerical Studies of Fundamental Flame Properties of Gaseous and Liquid Fuels

    DTIC Science & Technology

    2006-12-01

    PROPERTIES OF GASEOUS AND LIQUID FUELS FA9550-04-1-0006 5C. PROGRAM ELEMENT NUMBER 61102F 6. AUTHOR(S) 5d. PROJECT NUMBER 2308 Se. TASK NUMBER FOKION N...gaseous and liquid fuels, including jet fuels and their surrogates, were considered. Flame ignition and extinction limits were determined experimentally...EXPERIMENTAL AND DETAILED NUMERICAL STUDIES OF FUNDAMENTAL FLAME PROPERTIES OF GASEOUS AND LIQUID FUELS (AFOSR Grant FA9550-04-1-0006) (Period: 11/1/2003

  13. Combustion characteristics of hydrogen. Carbon monoxide based gaseous fuels

    NASA Technical Reports Server (NTRS)

    Notardonato, J. J.; White, D. J.; Kubasco, A. J.; Lecren, R. T.

    1981-01-01

    An experimental rig program was conducted with the objective of evaluating the combuston performance of a family of fuel gases based on a mixture of hydrogen and carbon monoxide. These gases, in addition to being members of a family, were also representative of those secondary fuels that could be produced from coal by various gasification schemes. In particular, simulated Winkler, Lurgi, and Blue-water low and medium energy content gases were used as fuels in the experimental combustor rig. The combustor used was originally designed as a low NOx rich-lean system for burning liquid fuels with high bound nitrogen levels. When used with the above gaseous fuels this combustor was operated in a lean-lean mode with ultra long residence times. The Blue-water gas was also operated in a rich-lean mode. The results of these tests indicate the possibility of the existence of an 'optimum' gas turbine hydrogen - carbon monoxide based secondary fuel. Such a fuel would exhibit NOx and high efficiency over the entire engine operating range. It would also have sufficient stability range to allow normal light-off and engine acceleration. Solar Turbines Incorporated would like to emphasize that the results presented here have been obtained with experimental rig combustors. The technologies generated could, however, be utilized in future commercial gas turbines.

  14. Combustion characteristics of hydrogen. Carbon monoxide based gaseous fuels

    NASA Astrophysics Data System (ADS)

    Notardonato, J. J.; White, D. J.; Kubasco, A. J.; Lecren, R. T.

    1981-10-01

    An experimental rig program was conducted with the objective of evaluating the combuston performance of a family of fuel gases based on a mixture of hydrogen and carbon monoxide. These gases, in addition to being members of a family, were also representative of those secondary fuels that could be produced from coal by various gasification schemes. In particular, simulated Winkler, Lurgi, and Blue-water low and medium energy content gases were used as fuels in the experimental combustor rig. The combustor used was originally designed as a low NOx rich-lean system for burning liquid fuels with high bound nitrogen levels. When used with the above gaseous fuels this combustor was operated in a lean-lean mode with ultra long residence times. The Blue-water gas was also operated in a rich-lean mode. The results of these tests indicate the possibility of the existence of an 'optimum' gas turbine hydrogen - carbon monoxide based secondary fuel. Such a fuel would exhibit NOx and high efficiency over the entire engine operating range. It would also have sufficient stability range to allow normal light-off and engine acceleration. Solar Turbines Incorporated would like to emphasize that the results presented here have been obtained with experimental rig combustors. The technologies generated could, however, be utilized in future commercial gas turbines.

  15. Fuel/propellant mixing in an open-cycle gas core nuclear rocket engine

    NASA Astrophysics Data System (ADS)

    Guo, Xu; Wehrmeyer, Joseph A.

    1997-01-01

    A numerical investigation of the mixing of gaseous uranium and hydrogen inside an open-cycle gas core nuclear rocket engine (spherical geometry) is presented. The gaseous uranium fuel is injected near the centerline of the spherical engine cavity at a constant mass flow rate, and the hydrogen propellant is injected around the periphery of the engine at a five degree angle to the wall, at a constant mass flow rate. The main objective is to seek ways to minimize the mixing of uranium and hydrogen by choosing a suitable injector geometry for the mixing of light and heavy gas streams. Three different uranium inlet areas are presented, and also three different turbulent models (k-ɛ model, RNG k-V model, and RSM model) are investigated. The commercial CFD code, FLUENT, is used to model the flow field. Uranium mole fraction, axial mass flux, and radial mass flux contours are obtained.

  16. Dual-fuel, dual-mode rocket engine

    NASA Technical Reports Server (NTRS)

    Martin, James A. (Inventor)

    1989-01-01

    The invention relates to a dual fuel, dual mode rocket engine designed to improve the performance of earth-to-orbit vehicles. For any vehicle that operates from the earth's surface to earth orbit, it is advantageous to use two different fuels during its ascent. A high density impulse fuel, such as kerosene, is most efficient during the first half of the trajectory. A high specific impulse fuel, such as hydrogen, is most efficient during the second half of the trajectory. The invention allows both fuels to be used with a single rocket engine. It does so by adding a minimum number of state-of-the-art components to baseline single made rocket engines, and is therefore relatively easy to develop for near term applications. The novelty of this invention resides in the mixing of fuels before exhaust nozzle cooling. This allows all of the engine fuel to cool the exhaust nozzle, and allows the ratio of fuels used throughout the flight depend solely on performance requirements, not cooling requirements.

  17. Method for providing real-time control of a gaseous propellant rocket propulsion system

    NASA Technical Reports Server (NTRS)

    Morris, Brian G. (Inventor)

    1991-01-01

    The new and improved methods and apparatus disclosed provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants. Real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time management of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

  18. Cooperative Testing of Rocket Injectors That Use Gaseous Oxygen and Hydrogen

    NASA Technical Reports Server (NTRS)

    1995-01-01

    Gaseous oxygen and hydrogen propellants used in a special engine energy cycle called Full-Flow Staged Combustion are believed to significantly increase the lifetime of a rocket engine's pumps. The cycle can also reduce the operating temperatures of the engine. Improving the lifetime of the hardware reduces its overall maintenance and operations costs, and is critical to reducing costs for the joint NASA/industry Reusable Launch Vehicle (RLV). The work in this project will demonstrate the performance and lifetime of one-element and many-element combustors with gaseous O2/H2 injectors. This work supporting the RLV program is a cooperative venture of the NASA Lewis Research Center, the NASA Marshall Space Flight Center, Rocketdyne, and the Pennsylvania State University. Information about gas-gas rocket injector performance with O2/H2 is very limited. Because of this paucity of data, new testing is needed to improve the knowledge base for testing and designing new injectors for the RLV and to improve computer models that predict the combusting gas flows of new injector designs. Therefore, detailed observations and measurements of the combusting flow from many-element injectors in a rocket engine are being sought. These observations and measurements will be done with three different tools: schlieren photography, ultraviolet imaging, and Raman spectroscopy. The schlieren system will take photos of the density differences in combusting flow, the ultraviolet movies will determine the location of the hydroxyl (OH) radical in the combustion flow, and the Raman spectroscopic measurements will provide the combustion temperature and amount of water (H2O), hydrogen (H2), and oxygen (O2) in the combustor. Marshall is providing overall program management, design and computational fluid dynamics (CFD) analyses, as well as funding for the work at Penn State. An existing, windowed combustor and several injectors will be provided by Rocketdyne--two injectors for the initial screening

  19. Fuel Regression Rate Behavior of CAMUI Hybrid Rocket

    NASA Astrophysics Data System (ADS)

    Kaneko, Yudai; Itoh, Mitsunori; Kakikura, Akihito; Mori, Kazuhiro; Uejima, Kenta; Nakashima, Takuji; Wakita, Masashi; Totani, Tsuyoshi; Oshima, Nobuyuki; Nagata, Harunori

    A series of static firing tests was conducted to investigate the fuel regression characteristics of a Cascaded Multistage Impinging-jet (CAMUI) type hybrid rocket motor. A CAMUI type hybrid rocket uses the combination of liquid oxygen and a fuel grain made of polyethylene as a propellant. The collision distance divided by the port diameter, H/D, was varied to investigate the effect of the grain geometry on the fuel regression rate. As a result, the H/D geometry has little effect on the regression rate near the stagnation point, where the heat transfer coefficient is high. On the contrary, the fuel regression rate decreases near the circumference of the forward-end face and the backward-end face of fuel blocks. Besides the experimental approaches, a method of computational fluid dynamics clarified the heat transfer distribution on the grain surface with various H/D geometries. The calculation shows the decrease of the flow velocity due to the increase of H/D on the area where the fuel regression rate decreases with the increase of H/D. To estimate the exact fuel consumption, which is necessary to design a fuel grain, real-time measurement by an ultrasonic pulse-echo method was performed.

  20. Production of gaseous fuel by pyrolysis of municipal solid waste

    NASA Technical Reports Server (NTRS)

    Crane, T. H.; Ringer, H. N.; Bridges, D. W.

    1975-01-01

    Pilot plant tests were conducted on a simulated solid waste which was a mixture of shredded newspaper, wood waste, polyethylene plastics, crushed glass, steel turnings, and water. Tests were conducted at 1400 F in a lead-bath pyrolyser. Cold feed was deaerated by compression and was dropped onto a moving hearth of molten lead before being transported to a sealed storage container. About 80 percent of the feed's organic content was converted to gaseous products which contain over 90 percent of the potential waste energy; 12 percent was converted to water; and 8 percent remained as partially pyrolyzed char and tars. Nearly half of the carbon in the feed is converted to benzene, toluene and medium-quality fuel gas, a potential credit of over $25 per ton of solid waste. The system was shown to require minimal preprocessing and less sorting then other methods.

  1. Vacuum plasma spray applications on liquid fuel rocket engines

    NASA Technical Reports Server (NTRS)

    Mckechnie, T. N.; Zimmerman, F. R.; Bryant, M. A.

    1992-01-01

    The vacuum plasma spray process (VPS) has been developed by NASA and Rocketdyne for a variety of applications on liquid fuel rocket engines, including the Space Shuttle Main Engine. These applications encompass thermal barrier coatings which are thermal shock resistant for turbopump blades and nozzles; bond coatings for cryogenic titanium components; wear resistant coatings and materials; high conductivity copper, NaRloy-Z, combustion chamber liners, and structural nickel base material, Inconel 718, for nozzle and combustion chamber support jackets.

  2. Vacuum plasma spray applications on liquid fuel rocket engines

    NASA Astrophysics Data System (ADS)

    McKechnie, T. N.; Zimmerman, F. R.; Bryant, M. A.

    1992-07-01

    The vacuum plasma spray process (VPS) has been developed by NASA and Rocketdyne for a variety of applications on liquid fuel rocket engines, including the Space Shuttle Main Engine. These applications encompass thermal barrier coatings which are thermal shock resistant for turbopump blades and nozzles; bond coatings for cryogenic titanium components; wear resistant coatings and materials; high conductivity copper, NaRloy-Z, combustion chamber liners, and structural nickel base material, Inconel 718, for nozzle and combustion chamber support jackets.

  3. Modeling of gaseous flows within proton exchange membrane fuel cells

    SciTech Connect

    Weisbrod, K.R.; Vanderborgh, N.E.; Grot, S.A.

    1996-12-31

    Development of a comprehensive mechanistic model has been helpful to understand PEM fuel cell performance. Both through-the-electrode and down-the-channel models have been developed to support our experimental effort to enhance fuel cell design and operation. The through-the-electrode model was described previously. This code describes the known transport properties and dynamic processes that occur within a membrane and electrode assembly. Key parameters include transport through the backing layers, water diffusion and electroosmotic transport in the membrane, and reaction electrochemical kinetics within the cathode catalyst layer. In addition, two geometric regions within the cathode layer are represented, the first region below saturation and second with liquid water present. Although processes at high gas stoichiometry are well represented by more simple codes, moderate stoichiometry processes require a two dimensional representation that include the gaseous composition and temperature along flow channel. Although usually PEM hardware utilizes serpentine flow channels, this code does not include such geometric features and thus the flow can be visualized along a single channel.

  4. Deposit formation and heat transfer in hydrocarbon rocket fuels

    NASA Technical Reports Server (NTRS)

    Giovanetti, A. J.; Spadaccini, L. J.; Szetela, E. J.

    1983-01-01

    An experimental research program was undertaken to investigate the thermal stability and heat transfer characteristics of several hydrocarbon fuels under conditions that simulate high-pressure, rocket engine cooling systems. The rates of carbon deposition in heated copper and nickel-plated copper tubes were determined for RP-1, propane, and natural gas using a continuous flow test apparatus which permitted independent variation and evaluation of the effect on deposit formation of wall temperature, fuel pressure, and fuel velocity. In addition, the effects of fuel additives and contaminants, cryogenic fuel temperatures, and extended duration testing with intermittent operation were examined. Parametric tests to map the thermal stability characteristics of RP-1, commercial-grade propane, and natural gas were conducted at pressures of 6.9 to 13.8 MPa, bulk fuel velocities of 30 to 90 m/s, and tube wall temperatures in the range of 230 to 810 K. Also, tests were run in which propane and natural gas fuels were chilled to 230 and 160 K, respectively. Corrosion of the copper tube surface was detected for all fuels tested. Plating the inside of the copper tubes with nickel reduced deposit formation and eliminated tube corrosion in most cases. The lowest rates of carbon deposition were obtained for natural gas, and the highest rates were obtained for propane. For all fuels tested, the forced-convection heat transfer film coefficients were satisfactorily correlated using a Nusselt-Reynolds-Prandtl number equation.

  5. Breakdown voltage determination of gaseous and near cryogenic fluids with application to rocket engine ignition

    NASA Astrophysics Data System (ADS)

    Nugent, Nicholas Jeremy

    density, spark gap distance, electrode angles, electrode materials and polarity. The research added to the fundamental knowledge of spark development in rocket ignition applications by determining the parameters that most influence breakdown voltage. Some improvements to the research should include better temperature measurements near the spark gap, additional testing with oxygen and testing with fuels of interest such as hydrogen and methane.

  6. Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

    NASA Technical Reports Server (NTRS)

    Bradley, David E.; Mireles, Omar R.; Hickman, Robert R.

    2011-01-01

    Deep space missions with large payloads require high specific impulse (Isp) and relatively high thrust in order to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average Isp. Nuclear thermal rockets (NTR) capable of high Isp thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high temperature hydrogen exposure on fuel elements is limited. The primary concern is the mechanical failure of fuel elements which employ high-melting-point metals, ceramics or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via non-contact RF heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

  7. Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

    NASA Technical Reports Server (NTRS)

    Bradley, D. E.; Mireles, O. R.; Hickman, R. R.

    2011-01-01

    Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames.1,2 Conventional storable propellants produce average specific impulse. Nuclear thermal rockets capable of producing high specific impulse are proposed. Nuclear thermal rockets employ heat produced by fission reaction to heat and therefore accelerate hydrogen, which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K), and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited.3 The primary concern is the mechanical failure of fuel elements that employ high-melting-point metals, ceramics, or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. The purpose of the testing is to obtain data to assess the properties of the non-nuclear support materials, as-fabricated, and determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures. The fission process of the planned fissile material and the resulting heating performance is well known and does not therefore require that active fissile material be integrated in this testing. A small-scale test bed designed to heat fuel element samples via non-contact radio frequency heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

  8. Fuel-Cell Power Source Based on Onboard Rocket Propellants

    NASA Technical Reports Server (NTRS)

    Ganapathi, Gani; Narayan, Sri

    2010-01-01

    The use of onboard rocket propellants (dense liquids at room temperature) in place of conventional cryogenic fuel-cell reactants (hydrogen and oxygen) eliminates the mass penalties associated with cryocooling and boil-off. The high energy content and density of the rocket propellants will also require no additional chemical processing. For a 30-day mission on the Moon that requires a continuous 100 watts of power, the reactant mass and volume would be reduced by 15 and 50 percent, respectively, even without accounting for boiloff losses. The savings increase further with increasing transit times. A high-temperature, solid oxide, electrolyte-based fuel-cell configuration, that can rapidly combine rocket propellants - both monopropellant system with hydrazine and bi-propellant systems such as monomethyl hydrazine/ unsymmetrical dimethyl hydrazine (MMH/UDMH) and nitrogen tetroxide (NTO) to produce electrical energy - overcomes the severe drawbacks of earlier attempts in 1963-1967 of using fuel reforming and aqueous media. The electrical energy available from such a fuel cell operating at 60-percent efficiency is estimated to be 1,500 Wh/kg of reactants. The proposed use of zirconia-based oxide electrolyte at 800-1,000 C will permit continuous operation, very high power densities, and substantially increased efficiency of conversion over any of the earlier attempts. The solid oxide fuel cell is also tolerant to a wide range of environmental temperatures. Such a system is built for easy refueling for exploration missions and for the ability to turn on after several years of transit. Specific examples of future missions are in-situ landers on Europa and Titan that will face extreme radiation and temperature environments, flyby missions to Saturn, and landed missions on the Moon with 14 day/night cycles.

  9. An experimental investigation of gaseous propellant rocket injectors using Raman spectroscopy

    NASA Astrophysics Data System (ADS)

    Foust, Michael Jerome

    The gaseous oxygen/gaseous hydrogen combusting flowfields of both shear and swirl coaxial injectors were studied in an optically-accessible uni-element rocket chamber. A Raman spectroscopy system was developed and applied for making spatially and temporally resolved measurements of the major combustion species in order to define the mixing and combustion characteristics of these injectors. Quantitative average species mole fraction and temperature distributions were obtained in the reacting flowfields at various axial locations downstream from both the shear and swirl coaxial elements. In addition to the Raman spectroscopy studies, the rate of heat transfer to the shear coaxial injector post from the flame was experimentally determined by obtaining temperature measurements within the injector post wall using thermocouples. The studies were conducted at three nominal chamber pressure conditions corresponding to 1.28 MPa, 2.12 MPa and 6.9 MPa. For the lowest chamber pressure condition being studied, the experimental measurements obtained downstream of a shear coaxial injector were in reasonable agreement with results from a computational fluid dynamics (CFD) code that was developed by Professor Merkle's group at The Pennsylvania State University; although the model does not capture the same level of unsteadiness in the flowfield as observed in the present studies. The results of both the experimental and numerical studies show that the shear coaxial element has slow mixing characteristics for the present operating conditions. In the studies conducted in the near-injector flowfield region of the shear coaxial element, the experimental results indicate that the stabilized position of the oxygen/hydrogen diffusion flame on the injector post is controlled by the relative momentum between the propellants at injection. In comparison to the flowfield characteristics of the shear coaxial injector, a swirl coaxial element displays higher mixing and combustion characteristics

  10. Metallic hydrogen: The most powerful rocket fuel yet to exist

    NASA Astrophysics Data System (ADS)

    Silvera, Isaac F.; Cole, John W.

    2010-03-01

    Wigner and Huntington first predicted that pressures of order 25 GPa were required for the transition of solid molecular hydrogen to the atomic metallic phase. Later it was predicted that metallic hydrogen might be a metastable material so that it remains metallic when pressure is released. Experimental pressures achieved on hydrogen have been more than an order of magnitude higher than the predicted transition pressure and yet it remains an insulator. We discuss the applications of metastable metallic hydrogen to rocketry. Metastable metallic hydrogen would be a very light-weight, low volume, powerful rocket propellant. One of the characteristics of a propellant is its specific impulse, Isp. Liquid (molecular) hydrogen-oxygen used in modern rockets has an Isp of ~460s; metallic hydrogen has a theoretical Isp of 1700s! Detailed analysis shows that such a fuel would allow single-stage rockets to enter into orbit or carry economical payloads to the moon. If pure metallic hydrogen is used as a propellant, the reaction chamber temperature is calculated to be greater than 6000 K, too high for currently known rocket engine materials. By diluting metallic hydrogen with liquid hydrogen or water, the reaction temperature can be reduced, yet there is still a significant performance improvement for the diluted mixture.

  11. Deposit formation and heat transfer in hydrocarbon rocket fuels

    NASA Technical Reports Server (NTRS)

    Giovanetti, A. J.; Spadaccini, L. J.; Szetela, E. J.

    1984-01-01

    An experimental research program was undertaken to investigate the thermal stability and heat transfer characteristics of several hydrocarbon fuels under conditions that simulate high-pressure, rocket engine cooling systems. The rates of carbon deposition in heated copper and nickel-plated copper tubes were determined for RP-1, propane, and natural gas using a continuous flow test apparatus which permitted independent variation and evaluation of the effect on deposit formation of wall temperature, fuel pressure, and fuel velocity. In addition, the effects of fuel additives and contaminants, cryogenic fuel temperatures, and extended duration testing with intermittent operation were examined. Corrosion of the copper tube surface was detected for all fuels tested; however, plating the insides of the tubes with nickel reduced deposit formation and eliminated corrosion in most cases. The lowest rates of carbon deposition were obtained for natural gas, and the highest rates were obtained for propane. Forced-convection heat transfer film coefficients were satisfactorily correlated using a Nusselt-Reynolds-Prandtl number equation for all the fuels tested.

  12. Method and system for low-NO.sub.x dual-fuel combustion of liquid and/or gaseous fuels

    SciTech Connect

    Gard, Vincent; Chojnacki, Dennis A; Rabovitser, Ioseph K

    2014-12-02

    A method and apparatus for combustion in which a pressurized preheated liquid fuel is atomized and a portion thereof flash vaporized, creating a mixture of fuel vapor and liquid droplets. The mixture is mixed with primary combustion oxidant, producing a fuel/primary oxidant mixture which is then injected into a primary combustion chamber in which the fuel/primary oxidant mixture is partially combusted, producing a secondary gaseous fuel containing hydrogen and carbon oxides. The secondary gaseous fuel is mixed with a secondary combustion oxidant and injected into the second combustion chamber wherein complete combustion of the secondary gaseous fuel is carried out. The resulting second stage flue gas containing very low amounts of NO.sub.x is then vented from the second combustion chamber.

  13. Moving, Moving, Moving- A Giant Rocket Fuel Tank

    NASA Image and Video Library

    2016-10-07

    Technicians moved a giant fuel tank from the Vertical Assembly Center where the tank recently completed friction stir welding to an adjacent work area at NASA's Michoud Assembly Facility in New Orleans. More than 1.7 miles of welds have been completed for core stage hardware at Michoud. This liquid hydrogen fuel tank is the largest piece of the core stage that will provide the fuel for the first flight of NASA's new rocket, the Space Launch System, with the Orion spacecraft in 2018. The tank is more than 130 feet long, and together with the liquid oxygen tank holds 733,000 gallons of propellant to feed the vehicle's four RS-25 engines to produce a total of 2 million pounds of thrust. SLS will have the power and capacity to carry humans to Mars. For more information on the core stage: http://www.nasa.gov/exploration/syste... Video Credit: NASA/MAF/Eric Bordelon

  14. Frequency-domain analysis for pulsating combustion of gaseous fuel

    NASA Astrophysics Data System (ADS)

    Berg, I. A.; Porshnev, S. V.; Oshchepkova, V. Y.; Medvedev, A. N.

    2017-06-01

    Pulsating combustion is among combustion control methods used to suppress formation of NOx. Past experiments showed that the dependency of NOx content from pulsation rate has a minimum. A measuring unit was set up to study torch behavior in infrared band. To study pulsating combustion of gaseous fuel a thermographic camera was used. Thermographic sequences were recorded using the instrument FLIR 7700M with the resolution of 320×240 pixels at the frame rate of 412 Hz. The experiments resulted in obtaining thermographic sequences radiation intensity fields in the longitudinal section of the torch at different pulsation rates. The obtained raw data was preprocessed to obtain distributions of quantities of pixels corresponding to temperatures in each frame, as well as time-domain series for changes of the torch core longitudinal section area. Frequency-domain analysis was run for each time-domain series using Fast Fourier transform (FFT). The results demonstrate that the first maximum of spectral density coincides with the control action rate. The spectrum also contains pronounced second and third harmonics. For each spectrum of the time-domain series signal-to-noise ratio (SNR) was calculated. Comparison of different SNR shows that maximum impact of pulsation control on torch radiation intensity takes place at the on/off valve opening rate of 4 Hz. This method of torch diagnostics can be helpful for future studies and development of pulsating combustion control systems.

  15. Stratified charge combustion system and method for gaseous fuel internal combustion engines

    SciTech Connect

    Rhoades, W.A. Jr.

    1986-03-11

    This patent describes a stratified charge combustion system for use in a gaseous fuel internal combustion engine. This system consists of: (a) a combustion chamber; (b) an ignition; (c) a gaseous fuel injection valve assembly in communication with the combustion chamber and in spaced relationship from the ignition source with a portion of the inside surfaces extending between the fuel injection valve assembly and the ignition source. The fuel valve assembly defines an entry port for the entrance of gaseous fuel, the entry port is recessed outside of a fixed inside surface. (d) means for pressuring the gaseous fuel prior to injection; and (e) a curved transitional surface extending from the entry port toward the portion of the inside surfaces extending between the fuel injection valve assembly and the ignition source. The curved transitional surface curves away from the direction of the entry port. The curved transitional surface has a curvature for the particular direction and configuration of the entry port. The particular configuration of the portion of the inside surfaces extends between the injection valve assembly and the ignition source. The particular arrangment of the fuel injection valve assembly in the combustion chamber, and for the particular pressure of the gaseous fuel is to produce the Coanda Effect in the injected gaseous fuel flow after it passes through the entry port and follows the curved transitional surface under the Coanda Effect. As the curved transitional surface curves away from the direction of the entry port, a flow is produced of the gaseous fuel that clings to and follows the particular configuration of the inside surfaces to the ignition source.

  16. Stability evaluation of a rocket engine for gaseous oxygen difluoride (OF2) and gaseous diborane (B2H6) propellants

    NASA Technical Reports Server (NTRS)

    Clayton, R. M.

    1972-01-01

    Results of an experimental evaluation of the dynamic stability of a candidate combustor for the space storable propellants gaseous OF2/B2H6 show that the combustor is unstable without supplementary damping. A computer analysis indicated that the uninhibited engine could be unstable. The experiments, conducted with O2/C2H4 substitute propellants and with 70-30 FLOX/B2H6 (OF2 simulated with FLOX), show that the uninhibited combustor has a low stability margin to starting transient perturbations, but that is relatively insensitive to bomb disturbances. Damping cavities are shown to provide stability.

  17. Solid Rocket Fuel Constitutive Theory and Polymer Cure

    NASA Technical Reports Server (NTRS)

    Ream, Robert

    2006-01-01

    Solid Rocket Fuel is a complex composite material for which no general constitutive theory, based on first principles, has been developed. One of the principles such a relation would depend on is the morphology of the binder. A theory of polymer curing is required to determine this morphology. During work on such a theory an algorithm was developed for counting the number of ways a polymer chain could assemble. The methods used to develop and check this algorithm led to an analytic solution to the problem. This solution is used in a probability distribution function which characterizes the morphology of the polymer.

  18. An americium-fueled gas core nuclear rocket

    SciTech Connect

    Kammash, T.; Galbraith, D.L.; Jan, T. )

    1993-01-10

    A gas core fission reactor that utilizes americium in place of uranium is examined for potential utilization as a nuclear rocket for space propulsion. The isomer [sup 242m]Am with a half life of 141 years is obtained from an (n, [gamma]) capture reaction with [sup 241]Am, and has the highest known thermal fission cross section. We consider a 7500 MW reactor, whose propulsion characteristics with [sup 235]U have already been established, and re-examine it using americium. We find that the same performance can be achieved at a comparable fuel density, and a radial size reduction (of both core and moderator/reflector) of about 70%.

  19. High Energy Density Additives for Hybrid Fuel Rockets to Improve Performance and Enhance Safety

    NASA Technical Reports Server (NTRS)

    Jaffe, Richard L.

    2014-01-01

    We propose a conceptual study of prototype strained hydrocarbon molecules as high energy density additives for hybrid rocket fuels to boost the performance of these rockets without compromising safety and reliability. Use of these additives could extend the range of applications for which hybrid rockets become an attractive alternative to conventional solid or liquid fuel rockets. The objectives of the study were to confirm and quantify the high enthalpy of these strained molecules and to assess improvement in rocket performance that would be expected if these additives were blended with conventional fuels. We confirmed the chemical properties (including enthalpy) of these additives. However, the predicted improvement in rocket performance was too small to make this a useful strategy for boosting hybrid rocket performance.

  20. The Gaseous Explosive Reaction : A Study of the Kinetics of Composite Fuels

    NASA Technical Reports Server (NTRS)

    Stevens, F W

    1929-01-01

    This report deals with the results of a series of studies of the kinetics of gaseous explosive reactions where the fuel under observation, instead of being a simple gas, is a known mixture of simple gases. In the practical application of the gaseous explosive reaction as a source of power in the gas engine, the fuels employed are composite, with characteristics that are apt to be due to the characteristics of their components and hence may be somewhat complex. The simplest problem that could be proposed in an investigation either of the thermodynamics or kinetics of the gaseous explosive reaction of a composite fuel would seem to be a separate study of the reaction characteristics of each component of the fuel and then a study of the reaction characteristics of the various known mixtures of those components forming composite fuels more and more complex. (author)

  1. Rocket Ignition Demonstrations Using Silane

    NASA Technical Reports Server (NTRS)

    Pal, Sibtosh; Santoro, Robert; Watkins, William B.; Kincaid, Kevin

    1998-01-01

    Rocket ignition demonstration tests using silane were performed at the Penn State Combustion Research Laboratory. A heat sink combustor with one injection element was used with gaseous propellants. Mixtures of silane and hydrogen were used as fuel, and oxygen was used as oxidizer. Reliable ignition was demonstrated using fuel lead and and a swirl injection element.

  2. Computational Thermochemistry of Jet Fuels and Rocket Propellants

    NASA Technical Reports Server (NTRS)

    Crawford, T. Daniel

    2002-01-01

    The design of new high-energy density molecules as candidates for jet and rocket fuels is an important goal of modern chemical thermodynamics. The NASA Glenn Research Center is home to a database of thermodynamic data for over 2000 compounds related to this goal, in the form of least-squares fits of heat capacities, enthalpies, and entropies as functions of temperature over the range of 300 - 6000 K. The chemical equilibrium with applications (CEA) program written and maintained by researchers at NASA Glenn over the last fifty years, makes use of this database for modeling the performance of potential rocket propellants. During its long history, the NASA Glenn database has been developed based on experimental results and data published in the scientific literature such as the standard JANAF tables. The recent development of efficient computational techniques based on quantum chemical methods provides an alternative source of information for expansion of such databases. For example, it is now possible to model dissociation or combustion reactions of small molecules to high accuracy using techniques such as coupled cluster theory or density functional theory. Unfortunately, the current applicability of reliable computational models is limited to relatively small molecules containing only around a dozen (non-hydrogen) atoms. We propose to extend the applicability of coupled cluster theory- often referred to as the 'gold standard' of quantum chemical methods- to molecules containing 30-50 non-hydrogen atoms. The centerpiece of this work is the concept of local correlation, in which the description of the electron interactions- known as electron correlation effects- are reduced to only their most important localized components. Such an advance has the potential to greatly expand the current reach of computational thermochemistry and thus to have a significant impact on the theoretical study of jet and rocket propellants.

  3. Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Larson, Daniel B.; Desain, John D.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth K.; Borduin, Russell; Koo, Joseph H.; Brady, Brian B.; Curtiss, Thomas J.; Story, George

    2012-01-01

    This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

  4. Liquid Oxygen Cooling of Hydrocarbon Fueled Rocket Thrust Chambers

    NASA Technical Reports Server (NTRS)

    Armstrong, Elizabeth S.

    1989-01-01

    Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the collant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

  5. Ablation study of tungsten-based nuclear thermal rocket fuel

    NASA Astrophysics Data System (ADS)

    Smith, Tabitha Elizabeth Rose

    The research described in this thesis has been performed in order to support the materials research and development efforts of NASA Marshall Space Flight Center (MSFC), of Tungsten-based Nuclear Thermal Rocket (NTR) fuel. The NTR was developed to a point of flight readiness nearly six decades ago and has been undergoing gradual modification and upgrading since then. Due to the simplicity in design of the NTR, and also in the modernization of the materials fabrication processes of nuclear fuel since the 1960's, the fuel of the NTR has been upgraded continuously. Tungsten-based fuel is of great interest to the NTR community, seeking to determine its advantages over the Carbide-based fuel of the previous NTR programs. The materials development and fabrication process contains failure testing, which is currently being conducted at MSFC in the form of heating the material externally and internally to replicate operation within the nuclear reactor of the NTR, such as with hot gas and RF coils. In order to expand on these efforts, experiments and computational studies of Tungsten and a Tungsten Zirconium Oxide sample provided by NASA have been conducted for this dissertation within a plasma arc-jet, meant to induce ablation on the material. Mathematical analysis was also conducted, for purposes of verifying experiments and making predictions. The computational method utilizes Anisimov's kinetic method of plasma ablation, including a thermal conduction parameter from the Chapman Enskog expansion of the Maxwell Boltzmann equations, and has been modified to include a tangential velocity component. Experimental data matches that of the computational data, in which plasma ablation at an angle shows nearly half the ablation of plasma ablation at no angle. Fuel failure analysis of two NASA samples post-testing was conducted, and suggestions have been made for future materials fabrication processes. These studies, including the computational kinetic model at an angle and the

  6. Comparison of thermal conversion methods of different biomass types into gaseous fuel

    NASA Astrophysics Data System (ADS)

    Larina, O. M.; Sinelshchikov, V. A.; Sytchev, G. A.

    2016-11-01

    Thermal conversion methods of different biomass types into gaseous fuel are considered. The comparison of the gas mixtures characteristics (volume yield, composition and calorific value) that can be produced from the main biomass types by gasification and pyrolysis is presented. The merits and demerits of these methods are discussed. It is shown that the two-stage pyrolysis technology, which consists of the biomass pyrolysis and the consequent high-temperature conversion of pyrolysis gases and vapors into synthesis gas by filtration through a porous carbon medium, allows to achieve both a high degree of biomass conversion into gaseous fuel and a high energy efficiency.

  7. Ultrasonic inspection of rocket fuel model using laminated transducer and multi-channel step pulser

    NASA Astrophysics Data System (ADS)

    Mihara, T.; Hamajima, T.; Tashiro, H.; Sato, A.

    2013-01-01

    For the ultrasonic inspection for the packing of solid fuel in a rocket booster, an industrial inspection is difficult. Because the signal to noise ratio in ultrasonic inspection of rocket fuel become worse due to the large attenuation even using lower frequency ultrasound. For the improvement of this problem, we tried to applied the two techniques in ultrasonic inspection, one was the step function pulser system with the super wideband frequency properties and the other was the laminated element transducer. By combining these two techniques, we developed the new ultrasonic measurement system and demonstrated the advantages in ultrasonic inspection of rocket fuel model specimen.

  8. Design and Testing of a Liquid Nitrous Oxide and Ethanol Fueled Rocket Engine

    SciTech Connect

    Youngblood, Stewart

    2015-08-01

    A small-scale, bi-propellant, liquid fueled rocket engine and supporting test infrastructure were designed and constructed at the Energetic Materials Research and Testing Center (EMRTC). This facility was used to evaluate liquid nitrous oxide and ethanol as potential rocket propellants. Thrust and pressure measurements along with high-speed digital imaging of the rocket exhaust plume were made. This experimental data was used for validation of a computational model developed of the rocket engine tested. The developed computational model was utilized to analyze rocket engine performance across a range of operating pressures, fuel-oxidizer mixture ratios, and outlet nozzle configurations. A comparative study of the modeling of a liquid rocket engine was performed using NASA CEA and Cantera, an opensource equilibrium code capable of being interfaced with MATLAB. One goal of this modeling was to demonstrate the ability of Cantera to accurately model the basic chemical equilibrium, thermodynamics, and transport properties for varied fuel and oxidizer operating conditions. Once validated for basic equilibrium, an expanded MATLAB code, referencing Cantera, was advanced beyond CEAs capabilities to predict rocket engine performance as a function of supplied propellant flow rate and rocket engine nozzle dimensions. Cantera was found to comparable favorably to CEA for making equilibrium calculations, supporting its use as an alternative to CEA. The developed rocket engine performs as predicted, demonstrating the developedMATLAB rocket engine model was successful in predicting real world rocket engine performance. Finally, nitrous oxide and ethanol were shown to perform well as rocket propellants, with specific impulses experimentally recorded in the range of 250 to 260 seconds.

  9. Modeling and Diagnostic Software for Liquefying-Fuel Rockets

    NASA Technical Reports Server (NTRS)

    Poll, Scott; Iverson, David; Ou, Jeremy; Sanderfer, Dwight; Patterson-Hine, Ann

    2005-01-01

    A report presents a study of five modeling and diagnostic computer programs considered for use in an integrated vehicle health management (IVHM) system during testing of liquefying-fuel hybrid rocket engines in the Hybrid Combustion Facility (HCF) at NASA Ames Research Center. Three of the programs -- TEAMS, L2, and RODON -- are model-based reasoning (or diagnostic) programs. The other two programs -- ICS and IMS -- do not attempt to isolate the causes of failures but can be used for detecting faults. In the study, qualitative models (in TEAMS and L2) and quantitative models (in RODON) having varying scope and completeness were created. Each of the models captured the structure and behavior of the HCF as a physical system. It was noted that in the cases of the qualitative models, the temporal aspects of the behavior of the HCF and the abstraction of sensor data are handled outside of the models, and it is necessary to develop additional code for this purpose. A need for additional code was also noted in the case of the quantitative model, though the amount of development effort needed was found to be less than that for the qualitative models.

  10. Rocket-Plume Spectroscopy Simulation for Hydrocarbon-Fueled Rocket Engines

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.

    2010-01-01

    The UV-Vis spectroscopic system for plume diagnostics monitors rocket engine health by using several analytical tools developed at Stennis Space Center (SSC), including the rocket plume spectroscopy simulation code (RPSSC), to identify and quantify the alloys from the metallic elements observed in engine plumes. Because the hydrocarbon-fueled rocket engine is likely to contain C2, CO, CH, CN, and NO in addition to OH and H2O, the relevant electronic bands of these molecules in the spectral range of 300 to 850 nm in the RPSSC have been included. SSC incorporated several enhancements and modifications to the original line-by-line spectral simulation computer program implemented for plume spectral data analysis and quantification in 1994. These changes made the program applicable to the Space Shuttle Main Engine (SSME) and the Diagnostic Testbed Facility Thruster (DTFT) exhaust plume spectral data. Modifications included updating the molecular and spectral parameters for OH, adding spectral parameter input files optimized for the 10 elements of interest in the spectral range from 320 to 430 nm and linking the output to graphing and analysis packages. Additionally, the ability to handle the non-uniform wavelength interval at which the spectral computations are made was added. This allowed a precise superposition of wavelengths at which the spectral measurements have been made with the wavelengths at which the spectral computations are done by using the line-by-line (LBL) code. To account for hydrocarbon combustion products in the plume, which might interfere with detection and quantification of metallic elements in the spectral region of 300 to 850 nm, the spectroscopic code has been enhanced to include the carbon-based combustion species of C2, CO, and CH. In addition, CN and NO have spectral bands in 300 to 850 nm and, while these molecules are not direct products of hydrocarbon-oxygen combustion systems, they can show up if nitrogen or a nitrogen compound is present

  11. High regression rate hybrid rocket fuel grains with helical port structures

    NASA Astrophysics Data System (ADS)

    Walker, Sean D.

    Hybrid rockets are popular in the aerospace industry due to their storage safety, simplicity, and controllability during rocket motor burn. However, they produce fuel regression rates typically 25% lower than solid fuel motors of the same thrust level. These lowered regression rates produce unacceptably high oxidizer-to-fuel (O/F) ratios that produce a potential for motor instability, nozzle erosion, and reduced motor duty cycles. To achieve O/F ratios that produce acceptable combustion characteristics, traditional cylindrical fuel ports are fabricated with very long length-to-diameter ratios to increase the total burning area. These high aspect ratios produce further reduced fuel regression rate and thrust levels, poor volumetric efficiency, and a potential for lateral structural loading issues during high thrust burns. In place of traditional cylindrical fuel ports, it is proposed that by researching the effects of centrifugal flow patterns introduced by embedded helical fuel port structures, a significant increase in fuel regression rates can be observed. The benefits of increasing volumetric efficiencies by lengthening the internal flow path will also be observed. The mechanisms of this increased fuel regression rate are driven by enhancing surface skin friction and reducing the effect of boundary layer "blowing" to enhance convective heat transfer to the fuel surface. Preliminary results using additive manufacturing to fabricate hybrid rocket fuel grains from acrylonitrile-butadiene-styrene (ABS) with embedded helical fuel port structures have been obtained, with burn-rate amplifications up to 3.0x than that of cylindrical fuel ports.

  12. Analysis of magma-thermal conversion of biomass to gaseous fuel

    SciTech Connect

    Gerlach, T.M.

    1982-02-01

    A wide range of magma types and pluton geometries believed to occur within the upper 10 km of the crust provide suitable sources of thermal energy for conversion of water-biomass mixtures to higher quality gaseous fuel. Gaseous fuel can be generated within a magma body, within the hot subsolidus margins of a magma body, or within surface reaction vessels heated by thermal energy derived from a magma body. The composition, amount, and energy content of the fuel gases generated from water-biomass mixtures are not sensitive to the type, age, depth, or temperature of a magma body thermal source. The amount and energy content of the generated fuel is almost entirely a function of the proportion of biomass in the starting mixture. CH/sub 4/ is the main gas that can be generated in important quantities by magma thermal energy under most circumstances. CO is never an important fuel product, and H/sub 2/ generation is very limited. The rates at which gaseous fuels can be generated are strongly dependent on magma type. Fuel generation rates for basaltic magmas are at least 2 to 3 times those for andesitic magmas and 5 to 6 times those for rhyolitic magmas. The highest fuel generation rates, for any particular magma body, will be achieved at the lowest possible reaction vessel operating temperature that does not cause graphite deposition from the water-biomass starting mixture. The energy content of the biomass-derived fuels is considerably greater than that consumed in the generation and refinement process.

  13. Apparatus and method for operating internal combustion engines from variable mixtures of gaseous fuels

    DOEpatents

    Heffel, James W.; Scott, Paul B.

    2003-09-02

    An apparatus and method for utilizing any arbitrary mixture ratio of multiple fuel gases having differing combustion characteristics, such as natural gas and hydrogen gas, within an internal combustion engine. The gaseous fuel composition ratio is first sensed, such as by thermal conductivity, infrared signature, sound propagation speed, or equivalent mixture differentiation mechanisms and combinations thereof which are utilized as input(s) to a "multiple map" engine control module which modulates selected operating parameters of the engine, such as fuel injection and ignition timing, in response to the proportions of fuel gases available so that the engine operates correctly and at high efficiency irrespective of the gas mixture ratio being utilized. As a result, an engine configured according to the teachings of the present invention may be fueled from at least two different fuel sources without admixing constraints.

  14. Apparatus and method for operating internal combustion engines from variable mixtures of gaseous fuels

    DOEpatents

    Heffel, James W.; Scott, Paul B.; Park, Chan Seung

    2011-11-01

    An apparatus and method for utilizing any arbitrary mixture ratio of multiple fuel gases having differing combustion characteristics, such as natural gas and hydrogen gas, within an internal combustion engine. The gaseous fuel composition ratio is first sensed, such as by thermal conductivity, infrared signature, sound propagation speed, or equivalent mixture differentiation mechanisms and combinations thereof which are utilized as input(s) to a "multiple map" engine control module which modulates selected operating parameters of the engine, such as fuel injection and ignition timing, in response to the proportions of fuel gases available so that the engine operates correctly and at high efficiency irrespective of the gas mixture ratio being utilized. As a result, an engine configured according to the teachings of the present invention may be fueled from at least two different fuel sources without admixing constraints.

  15. Measuring the Effect of Fuel Chemical Structure on Particulate and Gaseous Emissions using Isotope Tracing

    SciTech Connect

    Buchholz, B A; Mueller, C J; Martin, G C; Upatnicks, A; Dibble, R W; Cheng, S

    2003-09-11

    Using accelerator mass spectrometry (AMS), a technique initially developed for radiocarbon dating and recently applied to internal combustion engines, carbon atoms within specific fuel molecules can be labeled and followed in particulate or gaseous emissions. In addition to examining the effect of fuel chemical structure on emissions, the specific source of carbon for PM can be identified if an isotope label exists in the appropriate fuel source. Existing work has focused on diesel engines, but the samples (soot collected on quartz filters or combustion gases captured in bombs or bags) are readily collected from large industrial combustors as well.

  16. Turbulence in a gaseous hydrogen-liquid oxygen rocket combustion chamber

    NASA Technical Reports Server (NTRS)

    Lebas, J.; Tou, P.; Ohara, J.

    1975-01-01

    The intensity of turbulence and the Lagrangian correlation coefficient for a LOX-GH2 rocket combustion chamber was determined from experimental measurements of tracer gas diffusion. A combination of Taylor's turbulent diffusion theory and a numerical method for solving the conservation equations of fluid mechanics was used to calculate these quantities. Taylor's theory was extended to consider the inhomogeneity of the turbulence field in the axial direction of the combustion chamber, and an exponential function was used to represent the Lagrangian correlation coefficient. The results indicate that the value of the intensity of turbulence reaches a maximum of 14% at a location about 7" downstream from the injector. The Lagrangian correlation coefficient associated with this value is given by the above exponential expression where alpha = 10,000/sec.

  17. Microwave Extraction of Lunar Water for Rocket Fuel

    NASA Technical Reports Server (NTRS)

    Ethridge, Edwin C.; Donahue, Benjamin; Kaukler, William

    2008-01-01

    Nearly 50% of the lunar surface is oxygen, present as oxides in silicate rocks and soil. Methods for reduction of these oxides could liberate the oxygen. Remote sensing has provided evidence of significant quantities of hydrogen possibly indicating hundreds of millions of metric tons, MT, of water at the lunar poles. If the presence of lunar water is verified, water is likely to be the first in situ resource exploited for human exploration and for LOX-H2 rocket fuel. In-Situ lunar resources offer unique advantages for space operations. Each unit of product produced on the lunar surface represents 6 units that need not to be launched into LEO. Previous studies have indicated the economic advantage of LOX for space tugs from LEO to GEO. Use of lunar derived LOX in a reusable lunar lander would greatly reduce the LEO mass required for a given payload to the moon. And Lunar LOX transported to L2 has unique advantages for a Mars mission. Several methods exist for extraction of oxygen from the soil. But, extraction of lunar water has several significant advantages. Microwave heating of lunar permafrost has additional important advantages for water extraction. Microwaves penetrate and heat from within not just at the surface and excavation is not required. Proof of concept experiments using a moon in a bottle concept have demonstrated that microwave processing of cryogenic lunar permafrost simulant in a vacuum rapidly and efficiently extracts water by sublimation. A prototype lunar water extraction rover was built and tested for heating of simulant. Microwave power was very efficiently delivered into a simulated lunar soil. Microwave dielectric properties (complex electric permittivity and magnetic permeability) of lunar regolith simulant, JSC-1A, were measured down to cryogenic temperatures and above room temperature. The microwave penetration has been correlated with the measured dielectric properties. Since the microwave penetration depth is a function of temperature

  18. Microwave Extraction of Lunar Water for Rocket Fuel

    NASA Technical Reports Server (NTRS)

    Ethridge, Edwin C.; Donahue, Benjamin; Kaukler, William

    2008-01-01

    Nearly 50% of the lunar surface is oxygen, present as oxides in silicate rocks and soil. Methods for reduction of these oxides could liberate the oxygen. Remote sensing has provided evidence of significant quantities of hydrogen possibly indicating hundreds of millions of metric tons, MT, of water at the lunar poles. If the presence of lunar water is verified, water is likely to be the first in situ resource exploited for human exploration and for LOX-H2 rocket fuel. In-Situ lunar resources offer unique advantages for space operations. Each unit of product produced on the lunar surface represents 6 units that need not to be launched into LEO. Previous studies have indicated the economic advantage of LOX for space tugs from LEO to GEO. Use of lunar derived LOX in a reusable lunar lander would greatly reduce the LEO mass required for a given payload to the moon. And Lunar LOX transported to L2 has unique advantages for a Mars mission. Several methods exist for extraction of oxygen from the soil. But, extraction of lunar water has several significant advantages. Microwave heating of lunar permafrost has additional important advantages for water extraction. Microwaves penetrate and heat from within not just at the surface and excavation is not required. Proof of concept experiments using a moon in a bottle concept have demonstrated that microwave processing of cryogenic lunar permafrost simulant in a vacuum rapidly and efficiently extracts water by sublimation. A prototype lunar water extraction rover was built and tested for heating of simulant. Microwave power was very efficiently delivered into a simulated lunar soil. Microwave dielectric properties (complex electric permittivity and magnetic permeability) of lunar regolith simulant, JSC-1A, were measured down to cryogenic temperatures and above room temperature. The microwave penetration has been correlated with the measured dielectric properties. Since the microwave penetration depth is a function of temperature

  19. Computed tomography measurement of gaseous fuel concentration by infrared laser light absorption

    NASA Astrophysics Data System (ADS)

    Kawazoe, Hiromitsu; Inagaki, Kazuhisa; Emi, Y.; Yoshino, Fumio

    1997-11-01

    A system to measure gaseous hydrocarbon distributions was devised, which is based on IR light absorption by C-H stretch mode of vibration and computed tomography method. It is called IR-CT method in the paper. Affection of laser light power fluctuation was diminished by monitoring source light intensity by the second IR light detector. Calibration test for methane fuel was carried out to convert spatial data of line absorption coefficient into quantitative methane concentration. This system was applied to three flow fields. The first is methane flow with lifted flame which is generated by a gourd-shaped fuel nozzle. Feasibility of the IR-CT method was confirmed through the measurement. The second application is combustion field with diffusion flame. Calibration to determine absorptivity was undertaken, and measured line absorption coefficient was converted spatial fuel concentration using corresponding temperature data. The last case is modeled in cylinder gas flow of internal combustion engine, where gaseous methane was led to the intake valve in steady flow state. The fuel gas flow simulates behavior of gaseous gasoline which is evaporated at intake valve tulip. Computed tomography measurement of inner flow is essentially difficult because of existence of surrounding wall. In this experiment, IR laser beam was led to planed portion by IR light fiber. It is found that fuel convection by airflow takes great part in air-fuel mixture formation and the developed IR-CT system to measure fuel concentration is useful to analyze air-fuel mixture formation process and to develop new combustors.

  20. Design of a Premixed Gaseous Rocket Engine Injector for Ethylene and Oxygen

    DTIC Science & Technology

    2006-12-01

    97 INITIAL DISTRIBUTION LIST...Figure 4. Laminar Flame Velocities Versus Equivalence ratio for various fuel-air mixtures at standard conditions2...Injector at orifice size of 1/32”, mass flow of 0.0236 kg/sec and orifice size of 1/16”, mass flow 0.0267 kg/sec

  1. Fuel age impacts on gaseous fission product capture during separations

    SciTech Connect

    Jubin, Robert T.; Soelberg, Nicolas R.; Strachan, Denis M.; Ilas, G.

    2012-09-21

    As a result of fuel reprocessing, volatile radionuclides will be released from the facility stack if no processes are put in place to remove them. The radionuclides that are of concern in this document are 3H, 14C, 85Kr, and 129 Rosnick 2007 I. The question we attempt to answer is how efficient must this removal process be for each of these radionuclides? To answer this question, we examine the three regulations that may impact the degree to which these radionuclides must be reduced before process gases can be released from the facility. These regulations are 40 CFR 61 (EPA 2010a), 40 CFR 190(EPA 2010b), and 10 CFR 20 (NRC 2012), and they apply to the total radonuclide release and to the dose to a particular organ – the thyroid. Because these doses can be divided amongst all the radionuclides in different ways and even within the four radionuclides in question, several cases are studied. These cases consider for the four analyzed radionuclides inventories produced for three fuel types—pressurized water reactor uranium oxide (PWR UOX), pressurized water reactor mixed oxide (PWR MOX), and advanced high-temperature gascooled reactor (AHTGR)—several burnup values and time out of reactor extending to 200 y. Doses to the maximum exposed individual (MEI) are calculated with the EPA code CAP-88 ( , 1992). Two dose cases are considered. The first case, perhaps unrealistic, assumes that all of the allowable dose is assigned to the volatile radionuclides. In lieu of this, for the second case a value of 10% of the allowable dose is arbitrarily selected to be assigned to the volatile radionuclides. The required decontamination factors (DFs) are calculated for both of these cases, including the case for the thyroid dose for which 14C and 129I are the main contributors. However, for completeness, for one fuel type and burnup, additional cases are provided, allowing 25% and 50% of the allowable dose to be assigned to the volatile radionuclides. Because 3H and 85Kr have

  2. Gaseous-fuel nuclear reactor research for multimegawatt power in space

    NASA Technical Reports Server (NTRS)

    Thom, K.; Schneider, R. T.; Helmick, H. H.

    1977-01-01

    In the gaseous-fuel reactor concept, the fissile material is contained in a moderator-reflector cavity and exists in the form of a flowing gas or plasma separated from the cavity walls by means of fluid mechanical forces. Temperatures in excess of structural limitations are possible for low-specific-mass power and high-specific-impulse propulsion in space. Experiments have been conducted with a canister filled with enriched UF6 inserted into a beryllium-reflected cavity. A theoretically predicted critical mass of 6 kg was measured. The UF6 was also circulated through this cavity, demonstrating stable reactor operation with the fuel in motion. Because the flowing gaseous fuel can be continuously processed, the radioactive waste in this type of reactor can be kept small. Another potential of fissioning gases is the possibility of converting the kinetic energy of fission fragments directly into coherent electromagnetic radiation, the nuclear pumping of lasers. Numerous nuclear laser experiments indicate the possibility of transmitting power in space directly from fission energy. The estimated specific mass of a multimegawatt gaseous-fuel reactor power system is from 1 to 5 kg/kW while the companion laser-power receiver station would be much lower in specific mass.

  3. Gaseous-fuel nuclear reactor research for multimegawatt power in space

    NASA Technical Reports Server (NTRS)

    Thom, K.; Schneider, R. T.; Helmick, H. H.

    1977-01-01

    In the gaseous-fuel reactor concept, the fissile material is contained in a moderator-reflector cavity and exists in the form of a flowing gas or plasma separated from the cavity walls by means of fluid mechanical forces. Temperatures in excess of structural limitations are possible for low-specific-mass power and high-specific-impulse propulsion in space. Experiments have been conducted with a canister filled with enriched UF6 inserted into a beryllium-reflected cavity. A theoretically predicted critical mass of 6 kg was measured. The UF6 was also circulated through this cavity, demonstrating stable reactor operation with the fuel in motion. Because the flowing gaseous fuel can be continuously processed, the radioactive waste in this type of reactor can be kept small. Another potential of fissioning gases is the possibility of converting the kinetic energy of fission fragments directly into coherent electromagnetic radiation, the nuclear pumping of lasers. Numerous nuclear laser experiments indicate the possibility of transmitting power in space directly from fission energy. The estimated specific mass of a multimegawatt gaseous-fuel reactor power system is from 1 to 5 kg/kW while the companion laser-power receiver station would be much lower in specific mass.

  4. Gas turbine materials evaluation program utilizing coal derived gaseous fuel

    NASA Astrophysics Data System (ADS)

    Williams, M. L.; Yates, C. C.; Manning, G. B.; Peterson, R. R.

    1981-03-01

    A gas turbine materials evaluation test facility under the sponsorship of the U.S. Department of Energy is described. The objective of the mobile test facility is to obtain dynamic and static test data on the erosion/corrosion characteristics of materials exposed to the hot products of the combustion of coal-derived fuels. The engine being utilized for the tests is the WR 24-7 aircraft turbojet unit reconfigurated to burn coke oven gas. Approximately 100 hours of engine operating time have been logged to date.

  5. A computer program for thermal radiation from gaseous rocket exhuast plumes (GASRAD)

    NASA Technical Reports Server (NTRS)

    Reardon, J. E.; Lee, Y. C.

    1979-01-01

    A computer code is presented for predicting incident thermal radiation from defined plume gas properties in either axisymmetric or cylindrical coordinate systems. The radiation model is a statistical band model for exponential line strength distribution with Lorentz/Doppler line shapes for 5 gaseous species (H2O, CO2, CO, HCl and HF) and an appoximate (non-scattering) treatment of carbon particles. The Curtis-Godson approximation is used for inhomogeneous gases, but a subroutine is available for using Young's intuitive derivative method for H2O with Lorentz line shape and exponentially-tailed-inverse line strength distribution. The geometry model provides integration over a hemisphere with up to 6 individually oriented identical axisymmetric plumes, a single 3-D plume, Shading surfaces may be used in any of 7 shapes, and a conical limit may be defined for the plume to set individual line-of-signt limits. Intermediate coordinate systems may specified to simplify input of plumes and shading surfaces.

  6. Performance of a single fuel-vaporizing combustor with six injectors adapted for gaseous hydrogen

    NASA Technical Reports Server (NTRS)

    Wear, Jerrold D; Smith, Arthur L

    1955-01-01

    The combustor was operated over a range of inlet-air pressures from 5.3 to 24.0 inches of mercury absolute and inlet-air reference velocities from 60 to 100 feet per second. The combustion efficiencies obtained with the six configurations varied from about 65 to 95 percent for a combustor temperature-rise range of 200 degrees to 1400 degrees F. At a temperature rise of 1200 degrees F (near-rated engine conditions), the spread in efficiencies of the six configurations was about 5 percent. Efficiencies in the range of 65 to 85 percent were obtained at operating conditions beyond the burning range of conventional jet fuels. A fuel-injector configuration that fed only gaseous hydrogen fuel into the standard liquid-fuel-vaporizing tubes generally gave the highest efficiencies. This configuration minimizes the possibility of combustion in the fuel-vaporizing tubes and could be easily adapted to the full-scale engine combustor.

  7. Space shuttle with common fuel tank for liquid rocket booster and main engines (supertanker space shuttle)

    NASA Technical Reports Server (NTRS)

    Thorpe, Douglas G.

    1991-01-01

    An operation and schedule enhancement is shown that replaces the four-body cluster (Space Shuttle Orbiter (SSO), external tank, and two solid rocket boosters) with a simpler two-body cluster (SSO and liquid rocket booster/external tank). At staging velocity, the booster unit (liquid-fueled booster engines and vehicle support structure) is jettisoned while the remaining SSO and supertank continues on to orbit. The simpler two-bodied cluster reduces the processing and stack time until SSO mate from 57 days (for the solid rocket booster) to 20 days (for the liquid rocket booster). The areas in which liquid booster systems are superior to solid rocket boosters are discussed. Alternative and future generation vehicles are reviewed to reveal greater performance and operations enhancements with more modifications to the current methods of propulsion design philosophy, e.g., combined cycle engines, and concentric propellant tanks.

  8. The potential of algae blooms to produce renewable gaseous fuel.

    PubMed

    Allen, E; Browne, J; Hynes, S; Murphy, J D

    2013-11-01

    Ulva lactuca (commonly known as sea letuce) is a green sea weed which dominates Green Tides or algae blooms. Green Tides are caused by excess nitrogen from agriculture and sewage outfalls resulting in eutrophication in shallow estuaries. Samples of U. lactuca were taken from the Argideen estuary in West Cork on two consecutive years. In year 1 a combination of three different processes/pretreatments were carried out on the Ulva. These include washing, wilting and drying. Biomethane potential (BMP) assays were carried out on the samples. Fresh Ulva has a biomethane yield of 183LCH4/kgVS. For dried, washed and macerated Ulva a BMP of 250LCH4/kgVS was achieved. The resource from the estuary in West Cork was shown to be sufficient to provide fuel to 264 cars on a year round basis. Mono-digestion of Ulva may be problematic; the C:N ratio is low and the sulphur content is high. In year 2 co-digestion trials with dairy slurry were carried out. These indicate a potential increase in biomethane output by 17% as compared to mono-digestion of Ulva and slurry. Copyright © 2013 Elsevier Ltd. All rights reserved.

  9. Fuel Chemistry and Combustion Distribution Effects on Rocket Engine Combustion Stability

    DTIC Science & Technology

    2012-01-25

    H2O2 into hot oxygen and steam, an oxidizer injector mounted on a shaft that can translate during the course of the test, a co-axial fuel injector...have higher rates of reaction. It is known that hydrogen , with a very high rate of reaction, tends to be a stable rocket fuel . This study seeks to... hydrogen , JP-8 & ethanol), and liquid fuels containing energetic additives (eg hydrides and aluminum) have been studied. The modeling is being used to

  10. Piezo-fluidic Gaseous Fuel MPI System for Natural Gas Fuelled IC Engines

    NASA Astrophysics Data System (ADS)

    Chen, Rui

    A fast response piezo-fluidic gaseous fuel injector system designed for natural gas fuelled internal combustion (IC) engines is described in this paper. The system consists mainly of no moving part fluidic gas injector and piezo controlling interface. It can be arranged as a multi-point injection (MPI) system for IC engine fuel control. Both steady state and dynamic characteristics were investigated on a laboratory test rig. A comprehensive jet attachment and switching simulation model was also developed and reported. The agreement between predicted and experimental results is shown to be good.

  11. Bioconversion of solar energy into gaseous fuel (biogas) at elevated temperatures

    NASA Astrophysics Data System (ADS)

    Pantskhava, E. S.

    Various bacterial systems that can be used to convert proteins, lipids, and polysaccharides into gaseous fuel (biogas) and thus to replace the diminishing sources of natural fuels are discussed. Consideration is given to the individual reaction stages (i.e., the hydrolytic, acidogenic, and methanogenic steps) of the anaerobic fermentation process that produces methane and CO2 from raw biochemicals and to particular bacterial cultures responsible for these reactions. Special attention is given to the effects of various conditions of culture maintenance, such as hydrogenation and the rate and the manner of substrate renewal, on the yield of CH4 in the industrial production of biogas.

  12. Liquefied Gaseous Fuels Safety and Environmental Control Assessment Program: second status report

    SciTech Connect

    1980-10-01

    This document is arranged in three volumes and reports on progress in the Liquefied Gaseous Fuels (LGF) Safety and Environmental Control Assessment Program made in fiscal Year (FY)-1979 and early FY-1980. Volume 3 contains reports from 6 government contractors on LPG, anhydrous ammonia, and hydrogen energy systems. Report subjects include: simultaneous boiling and spreading of liquefied petroleum gas (LPG) on water; LPG safety research; state-of-the-art of release prevention and control technology in the LPG industry; ammonia: an introductory assessment of safety and environmental control information; ammonia as a fuel, and hydrogen safety and environmental control assessment.

  13. On-Line Measurement of Heat of Combustion of Gaseous Hydrocarbon Fuel Mixtures

    NASA Technical Reports Server (NTRS)

    Sprinkle, Danny R.; Chaturvedi, Sushil K.; Kheireddine, Ali

    1996-01-01

    A method for the on-line measurement of the heat of combustion of gaseous hydrocarbon fuel mixtures has been developed and tested. The method involves combustion of a test gas with a measured quantity of air to achieve a preset concentration of oxygen in the combustion products. This method involves using a controller which maintains the fuel (gas) volumetric flow rate at a level consistent with the desired oxygen concentration in the combustion products. The heat of combustion is determined form a known correlation with the fuel flow rate. An on-line computer accesses the fuel flow data and displays the heat of combustion measurement at desired time intervals. This technique appears to be especially applicable for measuring heats of combustion of hydrocarbon mixtures of unknown composition such as natural gas.

  14. Gas separation process using membranes with permeate sweep to remove CO.sub.2 from gaseous fuel combustion exhaust

    DOEpatents

    Wijmans, Johannes G [Menlo Park, CA; Merkel, Timothy C [Menlo Park, CA; Baker, Richard W [Palo Alto, CA

    2012-05-15

    A gas separation process for treating exhaust gases from the combustion of gaseous fuels, and gaseous fuel combustion processes including such gas separation. The invention involves routing a first portion of the exhaust stream to a carbon dioxide capture step, while simultaneously flowing a second portion of the exhaust gas stream across the feed side of a membrane, flowing a sweep gas stream, usually air, across the permeate side, then passing the permeate/sweep gas back to the combustor.

  15. Shaping of fuel delivery characteristics for solenoid operated diesel engine gaseous injectors

    SciTech Connect

    Hong, H.; Krepec, T.; Kekedjian, H.

    1996-09-01

    Solenoid operated gaseous injectors, when compared to conventional liquid fuel diesel injectors, differ in the way the fuel dose and its discharge rate are controlled. While in conventional diesel systems, the fuel dose and its injection rate depends on the fuel injection pump effective stroke and on the plunger diameter and velocity, the solenoid injectors operate in an on-off manner which limits the ability to control the gas discharge rate, resulting in its profile to be basically rectangular in shape. To reduce the gas injection rate at the beginning of the injection process in order to suppress the diesel-knock phenomenon, similar procedures as used in diesel engines could be implemented. One such approach is to use a throttling type pintle nozzle, and another method is to use a double-spring injector with a hole nozzle. The rationale for using such nozzle configurations is that gaseous fuels do not require atomization, and therefore, can be injected at lower discharge velocities than with liquid fuels. The gas delivery characteristics from a solenoid injector has been computer-simulated in order to assess the impact of the investigated three modes of fuel discharge rate control strategies. The simulation results confirmed that the gas dose and its discharge rate can be shaped as required. An experimental set-up is described to measure the gas discharge rate using a special gas injection mass flow rate indicator with a strain-gage sensor installed at the entry to a long tube, similar to that proposed by Bosch for liquid fuel volumetric flow rate measurements.

  16. High energy-density liquid rocket fuel performance

    NASA Technical Reports Server (NTRS)

    Rapp, Douglas C.

    1990-01-01

    A fuel performance database of liquid hydrocarbons and aluminum-hydrocarbon fuels was compiled using engine parametrics from the Space Transportation Engine Program as a baseline. Propellant performance parameters are introduced. General hydrocarbon fuel performance trends are discussed with respect to hydrogen-to-carbon ratio and heat of formation. Aluminum-hydrocarbon fuel performance is discussed with respect to aluminum metal loading. Hydrocarbon and aluminum-hydrocarbon fuel performance is presented with respect to fuel density, specific impulse, and propellant density specific impulse.

  17. High energy-density liquid rocket fuel performance

    NASA Technical Reports Server (NTRS)

    Rapp, Douglas C.

    1990-01-01

    A fuel performance database of liquid hydrocarbons and aluminum-hydrocarbon fuels was compiled using engine parametrics from the Space Transportation Engine Program as a baseline. Propellant performance parameters are introduced. General hydrocarbon fuel performance trends are discussed with respect to hydrogen-to-carbon ratio and heat of formation. Aluminum-hydrocarbon fuel performance is discussed with respect to aluminum metal loading. Hydrocarbon and aluminum-hydrocarbon fuel performance is presented with respect to fuel density, specific impulse and propellant density specific impulse.

  18. High energy-density liquid rocket fuel performance

    NASA Technical Reports Server (NTRS)

    Rapp, Douglas C.

    1990-01-01

    A fuel performance database of liquid hydrocarbons and aluminum-hydrocarbon fuels was compiled using engine parametrics from the Space Transportation Engine Program as a baseline. Propellant performance parameters are introduced. General hydrocarbon fuel performance trends are discussed with respect to hydrogen-to-carbon ratio and heat of formation. Aluminum-hydrocarbon fuel performance is discussed with respect to aluminum metal loading. Hydrocarbon and aluminum-hydrocarbon fuel performance is presented with respect to fuel density, specific impulse and propellant density specific impulse.

  19. Development of carbon-carbon nozzle extension for liquid fuel rocket motors

    NASA Astrophysics Data System (ADS)

    Sokolovsky, M. I.; Petukhov, S. N.; Semyonov, Yu. P.; Sokolov, B. A.

    2008-12-01

    Successful experience of RSC “Energy” and SPA “Iskra” in the development of carbon-carbon extension for oxygen-kerosene liquid fuel rocket motor has been summarized. Methodological approach that served to completion of carbon-carbon extension development in full and at comparatively small expenses has been described. Results of practical application of carbon-carbon extension for liquid fuel rocket motor 11D58M have been presented within the framework of International Space Program “Sea Launch”.

  20. Integrated model development for liquid fueled rocket propulsion systems

    NASA Technical Reports Server (NTRS)

    Santi, L. Michael

    1993-01-01

    As detailed in the original statement of work, the objective of phase two of this research effort was to develop a general framework for rocket engine performance prediction that integrates physical principles, a rigorous mathematical formalism, component level test data, system level test data, and theory-observation reconciliation. Specific phase two development tasks are defined.

  1. Solid amine-boranes as high performance hypergolic hybrid rocket fuels

    NASA Astrophysics Data System (ADS)

    Pfeil, Mark A.

    Hypergolic hybrid rockets have the potential of providing systems that are simple, reliable, have high performance, and allow for energy management. Such a propulsion system can be applied to fields that need a single tactical motor with flexible mission requirements of either high speed to target or extended loitering. They also provide the possibility for alternative fast response dynamic altitude control systems if ignition delays are sufficiently short. Amines are the traditional fuel of choice when selecting a hypergolic combination as these tend to react readily with both nitric acid and dinitrogen tertroxide based oxidizers. It has been found that the addition of a borane adduct to an amine fuel tends to reduce the ignition delay by up to an order of magnitude with white fuming nitric acid (WFNA). The borane addition has resulted in fuels with very short ignition delays between 2-10 ms - the fastest times for an amine based fuel reacting with nitric acid based oxidizers. The incorporation of these amine-boranes, specifically ethylenediamine bisborane (EDBB), into various fuel binders has also been found to result in ignition delays between 3-10 ms - the fastest times again for amine based fuels. It was found that the addition of a borane to an amine increased theoretical performance of the amine resulting in high performance fuels. The amine-borane/fuel binder combinations also produced higher theoretical performance values than previously used hypergolic hybrid rockets. Some of the theoretical values are on par or higher than the current toxic liquid hypergolic fuels, making amine boranes an attractive replacement. The higher performing amine-borane/fuel binder combinations also have higher performance values than the traditional rocket fuels, excluding liquid hydrogen. Thus, amine-borane based fuels have the potential to influence various area in the rocket field. An EDBB/ferrocene/epoxy fuel was tested in a hypergolic hybrid with pure nitric acid as the

  2. Pollutant Emissions and Lean Blowoff Limits of Fuel Flexible Burners Operating on Gaseous Renewable and Fossil Fuels

    NASA Astrophysics Data System (ADS)

    Colorado, Andres

    This study provides an experimental and numerical examination of pollutant emissions and stability of gaseous fueled reactions stabilized with two premixed-fuel-flexible and ultra-low NOx burner technologies. Both burners feature lean combustion technology to control the formation of nitrogen oxides (NOx). The first fuel--flexible burner is the low-swirl burner (LSB), which features aerodynamic stabilization of the reactions with a divergent flow-field; the second burner is the surface stabilized combustion burner (SSCB), which features the stabilization of the reactions on surface patterns. For combustion applications the most commonly studied species are: NOx, carbon monoxide (CO), and unburned hydrocarbons (UHC). However these are not the only pollutants emitted when burning fossil fuels; other species such as nitrous oxide (N2O), ammonia (NH3) and formaldehyde (CH2O) can be directly emitted from the oxidation reactions. Yet the conditions that favor the emission of these pollutants are not completely understood and require further insight. The results of this dissertation close the gap existing regarding the relations between emission of pollutants species and stability when burning variable gaseous fuels. The results of this study are applicable to current issues such as: 1. Current combustion systems operating at low temperatures to control formation of NOx. 2. Increased use of alternative fuels such as hydrogen, synthetic gas and biogas. 3. Increasing recognition of the need/desire to operate combustion systems in a transient manner to follow load and to offset the intermittency of renewable power. 4. The recent advances in measurement methods allow us to quantify other pollutants, such as N 2O, NH3 and CH2O. Hence in this study, these pollutant species are assessed when burning natural gas (NG) and its binary mixtures with other gaseous fuels such as hydrogen (H2), carbon dioxide (CO2), ethane (C 2H6) and propane (C3H8) at variable operation modes including

  3. Thermal hydraulic design analysis of ternary carbide fueled square-lattice honeycomb nuclear rocket engine

    SciTech Connect

    Furman, Eric M.; Anghaie, Samim

    1999-01-22

    A computational analysis is conducted to determine the optimum thermal-hydraulic design parameters for a square-lattice honeycomb nuclear rocket engine core that will incorporate ternary carbide based uranium fuels. Recent studies at the Innovative Nuclear Space Power and Propulsion Institute (INSPI) have demonstrated the feasibility of processing solid solution, ternary carbide fuels such as (U, Zr, Nb)C, (U, Zr, Ta)C, (U, Zr, Hf)C and (U, Zr, W)C. The square-lattice honeycomb design provides high strength and is amenable to the processing complexities of these ultrahigh temperature fuels. A parametric analysis is conducted to examine how core geometry, fuel thickness and the propellant flow area effect the thermal performance of the nuclear rocket engine. The principal variables include core size (length and diameter) and fuel element dimensions. The optimum core configuration requires a balance between high specific impulse and thrust level performance, and maintaining the temperature and strength limits of the fuel. A nuclear rocket engine simulation code is developed and used to examine the system performance as well as the performance of the main reactor core components. The system simulation code was originally developed for analysis of NERVA-Derivative and Pratt and Whitney XNR-2000 nuclear thermal rockets. The code is modified and adopted to the square-lattice geometry of the new fuel design. Thrust levels ranging from 44,500 to 222,400 N (10,000 to 50,000 lbf) are considered. The average hydrogen exit temperature is kept at 2800 K, which is well below the melting point of these fuels. For a nozzle area ratio of 300 and a thrust chamber pressure of 4.8 Mpa (700 psi), the specific impulse is 930 s. Hydrogen temperature and pressure distributions in the core and the fuel maximum temperatures are calculated.

  4. Thermal hydraulic design analysis of ternary carbide fueled square-lattice honeycomb nuclear rocket engine

    NASA Astrophysics Data System (ADS)

    Furman, Eric M.; Anghaie, Samim

    1999-01-01

    A computational analysis is conducted to determine the optimum thermal-hydraulic design parameters for a square-lattice honeycomb nuclear rocket engine core that will incorporate ternary carbide based uranium fuels. Recent studies at the Innovative Nuclear Space Power and Propulsion Institute (INSPI) have demonstrated the feasibility of processing solid solution, ternary carbide fuels such as (U, Zr, Nb)C, (U, Zr, Ta)C, (U, Zr, Hf)C and (U, Zr, W)C. The square-lattice honeycomb design provides high strength and is amenable to the processing complexities of these ultrahigh temperature fuels. A parametric analysis is conducted to examine how core geometry, fuel thickness and the propellant flow area effect the thermal performance of the nuclear rocket engine. The principal variables include core size (length and diameter) and fuel element dimensions. The optimum core configuration requires a balance between high specific impulse and thrust level performance, and maintaining the temperature and strength limits of the fuel. A nuclear rocket engine simulation code is developed and used to examine the system performance as well as the performance of the main reactor core components. The system simulation code was originally developed for analysis of NERVA-Derivative and Pratt & Whitney XNR-2000 nuclear thermal rockets. The code is modified and adopted to the square-lattice geometry of the new fuel design. Thrust levels ranging from 44,500 to 222,400 N (10,000 to 50,000 lbf) are considered. The average hydrogen exit temperature is kept at 2800 K, which is well below the melting point of these fuels. For a nozzle area ratio of 300 and a thrust chamber pressure of 4.8 Mpa (700 psi), the specific impulse is 930 s. Hydrogen temperature and pressure distributions in the core and the fuel maximum temperatures are calculated.

  5. Combustion of solid fuel slabs with gaseous oxygen in a hybrid motor analog

    NASA Technical Reports Server (NTRS)

    Chiaverini, Martin J.; Harting, George C.; Lu, Yeu-Cherng; Kuo, Kenneth K.; Serin, Nadir; Johnson, David K.

    1995-01-01

    Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated- Polybutadiene) fuel cross linked with diisocyanate was burned with GOX under various operating conditions. Large amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed line system and combustion chamber, the pressure oscillations were drastically reduced from +/- 20% of the localized mean pressure to an acceptable range of +/- 1.5%. Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations arc thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thicknesses and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented. Several tests were conducted using, simultaneously, one translucent fuel slab and one fuel slab processed with carbon black powder. The addition of carbon black did not affect the measured regression rates or surface temperatures in comparison

  6. Combustion of solid fuel slabs with gaseous oxygen in a hybrid motor analog

    NASA Technical Reports Server (NTRS)

    Chiaverini, Martin J.; Harting, George C.; Lu, Yeu-Cherng; Kuo, Kenneth K.; Serin, Nadir; Johnson, David K.

    1995-01-01

    Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated- Polybutadiene) fuel cross linked with diisocyanate was burned with GOX under various operating conditions. Large amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed line system and combustion chamber, the pressure oscillations were drastically reduced from +/- 20% of the localized mean pressure to an acceptable range of +/- 1.5%. Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations arc thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thicknesses and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented. Several tests were conducted using, simultaneously, one translucent fuel slab and one fuel slab processed with carbon black powder. The addition of carbon black did not affect the measured regression rates or surface temperatures in comparison

  7. High turbulence combustion chamber for turbocharged lean burn gaseous fueled engine

    SciTech Connect

    Joyce, R.S.

    1988-12-27

    This patent describes a gaseous fueled engine having at least a cylinder with an fuel/air intake and combustion product exhaust to be used in a cylinder head having a flat lower surface enclosing one end of the cylinder, and a piston with a flat top piston head slidably received in the cylinder. The improvement consists of: a. an ignition device extending from the flat lower surface of the cylinder head in the center of the cylinder, and b. a baffleless combustion chamber eccentrically located in the head of the piston with respect to the piston axis such that ignition of the fuel/air mixture by the ignition device occurs within but on the outer periphery of the combustion chamber.

  8. Gaseous Species Measurements of Alternative Jet Fuels in Sooting Laminar Coflow Diffusion Flames

    NASA Astrophysics Data System (ADS)

    Zabeti, Parham

    The gaseous species concentration of Jet A-1, GTL, CTL and a blend of 80 vol.% GTL and 20 vol.% hexanol jet fuels in laminar coflow diffusion flames have been measured and studied. These species are carbon monoxide, carbon dioxide, oxygen, methane, ethane, ethylene, propylene, and acetylene. Benzene and propyne concentrations were also detected in CTL flames. 1-Butene has been quantified for the blend of GTL and hexanol flame. The detailed experimental setup has been described and results from different flames are compared. The CO is produced in a same amount in all the flames. The CTL flame had the largest and GTL/hexanol flame had lowest CO2 concentrations. The results indicate that GTL and GTL hexanol blend flames produce similar concentrations for all the measured hydrocarbon species and have the highest concentration among all the jet fuels. The experimental results from Jet A-1 fuel are also compared with numerical studies by Saffaripour et al .

  9. Nonlinear longitudinal oscillations of fuel in rockets feed lines with gas-liquid damper

    NASA Astrophysics Data System (ADS)

    Avramov, K. V.; Filipkovsky, S.; Tonkonogenko, A. M.; Klimenko, D. V.

    2016-03-01

    The mathematical model of the fuel oscillations in the rockets feed lines with gas-liquid dampers is derived. The nonlinear model of the gas-liquid damper is suggested. The vibrations of fuel in the feed lines with the gas-liquid dampers are considered nonlinear. The weighted residual method is applied to obtain the finite degrees of freedom nonlinear model of the fuel oscillations. Shaw-Pierre nonlinear normal modes are applied to analyze free vibrations. The forced oscillations of the fuel at the principle resonances are analyzed. The stability of the forced oscillations is investigated. The results of the forced vibrations analysis are shown on the frequency responses.

  10. Hybrid rocket fuel combustion and regression rate study

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Ray, R. L.; Anderson, F. A.; Cohen, N. S.

    1992-01-01

    The objectives of this study are to develop hybrid fuels (1) with higher regression rates and reduced dependence on fuel grain geometry and (2) that maximize potential specific impulse using low-cost materials. A hybrid slab window motor system was developed to screen candidate fuels - their combustion behavior and regression rate. Combustion behavior diagnostics consisted of video and high speed motion pictures coverage. The mean fuel regression rates were determined by before and after measurements of the fuel slabs. The fuel for this initial investigation consisted of hydroxyl-terminated polybutadiene binder with coal and aluminum fillers. At low oxidizer flux levels (and corresponding fuel regression rates) the filled-binder fuels burn in a layered fashion, forming an aluminum containing binder/coal surface melt that, in turn, forms into filigrees or flakes that are stripped off by the crossflow. This melt process appears to diminish with increasing oxidizer flux level. Heat transfer by radiation is a significant contributor, producing the desired increase in magnitude and reduction in flow dependency (power law exponent) of the fuel regression rate.

  11. Hybrid rocket fuel combustion and regression rate study

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Ray, R. L.; Anderson, F. A.; Cohen, N. S.

    1992-01-01

    The objectives of this study are to develop hybrid fuels (1) with higher regression rates and reduced dependence on fuel grain geometry and (2) that maximize potential specific impulse using low-cost materials. A hybrid slab window motor system was developed to screen candidate fuels - their combustion behavior and regression rate. Combustion behavior diagnostics consisted of video and high speed motion pictures coverage. The mean fuel regression rates were determined by before and after measurements of the fuel slabs. The fuel for this initial investigation consisted of hydroxyl-terminated polybutadiene binder with coal and aluminum fillers. At low oxidizer flux levels (and corresponding fuel regression rates) the filled-binder fuels burn in a layered fashion, forming an aluminum containing binder/coal surface melt that, in turn, forms into filigrees or flakes that are stripped off by the crossflow. This melt process appears to diminish with increasing oxidizer flux level. Heat transfer by radiation is a significant contributor, producing the desired increase in magnitude and reduction in flow dependency (power law exponent) of the fuel regression rate.

  12. Model of lidar range-Doppler signatures of solid rocket fuel plumes

    NASA Astrophysics Data System (ADS)

    Bankman, Isaac N.; Giles, John W.; Chan, Stephen C.; Reed, Robert A.

    2004-09-01

    The analysis of particles produced by solid rocket motor fuels relates to two types of studies: the effect of these particles on the Earth's ozone layer, and the dynamic flight behavior of solid fuel boosters used by the NASA Space Shuttle. Since laser backscatter depends on the particle size and concentration, a lidar system can be used to analyze the particle distributions inside a solid rocket plume in flight. We present an analytical model that simulates the lidar returns from solid rocket plumes including effects of beam profile, spot size, polarization and sensing geometry. The backscatter and extinction coefficients of alumina particles are computed with the T-matrix method that can address non-spherical particles. The outputs of the model include time-resolved return pulses and range-Doppler signatures. Presented examples illustrate the effects of sensing geometry.

  13. Determination of Combustion Product Radicals in a Hydrocarbon Fueled Rocket Exhaust Plume

    NASA Technical Reports Server (NTRS)

    Langford, Lester A.; Allgood, Daniel C.; Junell, Justin C.

    2007-01-01

    The identification of metallic effluent materials in a rocket engine exhaust plume indicates the health of the engine. Since 1989, emission spectroscopy of the plume of the Space Shuttle Main Engine (SSME) has been used for ground testing at NASA's Stennis Space Center (SSC). This technique allows the identification and quantification of alloys from the metallic elements observed in the plume. With the prospect of hydrocarbon-fueled rocket engines, such as Rocket Propellant 1 (RP-1) or methane (CH4) fueled engines being considered for use in future space flight systems, the contributions of intermediate or final combustion products resulting from the hydrocarbon fuels are of great interest. The effect of several diatomic molecular radicals, such as Carbon Dioxide , Carbon Monoxide, Molecular Carbon, Methylene Radical, Cyanide or Cyano Radical, and Nitric Oxide, needs to be identified and the effects of their band systems on the spectral region from 300 nm to 850 nm determined. Hydrocarbon-fueled rocket engines will play a prominent role in future space exploration programs. Although hydrogen fuel provides for higher engine performance, hydrocarbon fuels are denser, safer to handle, and less costly. For hydrocarbon-fueled engines using RP-1 or CH4 , the plume is different from a hydrogen fueled engine due to the presence of several other species, such as CO2, C2, CO, CH, CN, and NO, in the exhaust plume, in addition to the standard H2O and OH. These species occur as intermediate or final combustion products or as a result of mixing of the hot plume with the atmosphere. Exhaust plume emission spectroscopy has emerged as a comprehensive non-intrusive sensing technology which can be applied to a wide variety of engine performance conditions with a high degree of sensitivity and specificity. Stennis Space Center researchers have been in the forefront of advancing experimental techniques and developing theoretical approaches in order to bring this technology to a more

  14. Exploiting hydrophobic borohydride-rich ionic liquids as faster-igniting rocket fuels.

    PubMed

    Liu, Tianlin; Qi, Xiujuan; Huang, Shi; Jiang, Linhai; Li, Jianling; Tang, Chenglong; Zhang, Qinghua

    2016-02-04

    A family of hydrophobic borohydride-rich ionic liquids was developed, which exhibited the shortest ignition delay times of 1.7 milliseconds and the lowest viscosity (10 mPa s) of hypergolic ionic fluids, demonstrating their great potential as faster-igniting rocket fuels to replace toxic hydrazine derivatives in liquid bipropellant formulations.

  15. Assessing the impacts of ethanol and isobutanol on gaseous and particulate emissions from flexible fuel vehicles.

    PubMed

    Karavalakis, Georgios; Short, Daniel; Russell, Robert L; Jung, Heejung; Johnson, Kent C; Asa-Awuku, Akua; Durbin, Thomas D

    2014-12-02

    This study investigated the effects of higher ethanol blends and an isobutanol blend on the criteria emissions, fuel economy, gaseous toxic pollutants, and particulate emissions from two flexible-fuel vehicles equipped with spark ignition engines, with one wall-guided direct injection and one port fuel injection configuration. Both vehicles were tested over triplicate Federal Test Procedure (FTP) and Unified Cycles (UC) using a chassis dynamometer. Emissions of nonmethane hydrocarbons (NMHC) and carbon monoxide (CO) showed some statistically significant reductions with higher alcohol fuels, while total hydrocarbons (THC) and nitrogen oxides (NOx) did not show strong fuel effects. Acetaldehyde emissions exhibited sharp increases with higher ethanol blends for both vehicles, whereas butyraldehyde emissions showed higher emissions for the butanol blend relative to the ethanol blends at a statistically significant level. Particulate matter (PM) mass, number, and soot mass emissions showed strong reductions with increasing alcohol content in gasoline. Particulate emissions were found to be clearly influenced by certain fuel parameters including oxygen content, hydrogen content, and aromatics content.

  16. Investigation of thermal and environmental characteristics of combustion of gaseous fuels

    NASA Astrophysics Data System (ADS)

    Vetkin, A. V.; Suris, A. L.

    2015-03-01

    Numerical investigations are fulfilled for some thermal and environmental characteristics of combustion of gaseous fuels used at present in tube furnaces of petroleum refineries. The effect of the fuel composition on these characteristics is shown and probable consequences of the substitution of natural gas to other types of fuels. Methane, ethane, propane, butane, propylene, and hydrogen are considered for comparison, which in most cases are constituents of the composition of the fuel burnt in furnaces. The effect of the fuel type, its associated combustion temperature, combustion product emissivity, temperature of combustion chamber walls, mean beam length, and heat release on the variation in the radiant heat flux within the radiant chamber of furnaces is investigated. The effect of flame characteristics, which are determined by the presence of diffusion combustion zones formed by burners used at present in furnaces for reducing nitrogen oxides emission, is analyzed. The effect of the fuel type on the equilibrium NO concentration is also investigated. The investigations were carried out both at arbitrary given gas temperatures and at effective temperatures dependent on the adiabatic combustion temperature and the temperature at the chamber output and determined based on solving a set of equations at various heat-release rates of the combustion chamber.

  17. 40 CFR Appendix Xvi to Part 86 - Pollutant Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles...

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 20 2012-07-01 2012-07-01 false Pollutant Mass Emissions Calculation... Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles Equipped With...-Fueled Vehicle Pollutant Mass Emission Calculation Procedure. (1) For all TLEVs, LEVs, and ULEVs, the...

  18. 40 CFR Appendix Xvi to Part 86 - Pollutant Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles...

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 19 2011-07-01 2011-07-01 false Pollutant Mass Emissions Calculation... Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles Equipped With...-Fueled Vehicle Pollutant Mass Emission Calculation Procedure. (1) For all TLEVs, LEVs, and ULEVs, the...

  19. 40 CFR Appendix Xvi to Part 86 - Pollutant Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles...

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 20 2013-07-01 2013-07-01 false Pollutant Mass Emissions Calculation... Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles Equipped With...-Fueled Vehicle Pollutant Mass Emission Calculation Procedure. (1) For all TLEVs, LEVs, and ULEVs, the...

  20. 40 CFR Appendix Xvi to Part 86 - Pollutant Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles...

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... National Low Emission Vehicle Program (October, 1996) shall apply. These procedures are incorporated by... Mass Emissions Calculation Procedure for Gaseous-Fueled Vehicles and for Vehicles Equipped With...-Fueled Vehicle Pollutant Mass Emission Calculation Procedure. (1) For all TLEVs, LEVs, and ULEVs,...

  1. The development of reactive fuel grains for pyrophoric relight of in-space hybrid rocket thrusters

    NASA Astrophysics Data System (ADS)

    Steiner, Matthew Wellington

    This study presents and investigates a novel hybrid fuel grain that reacts pyrophorically with gaseous oxidizer to achieve restart of a hybrid rocket motor propulsion system while reducing cost and handling concerns. This reactive fuel grain (RFG) relies on the pyrophoric nature of finely divided metal particles dispersed in a solid dicyclopentadiene (DCPD) binder, which has been shown to encapsulate air-sensitive additives until they are exposed to combustion gases. An RFG is thus effectively inert in open air in the absence of an ignition source, though the particles encapsulated within remain pyrophoric. In practice, this means that an RFG that is ignited in the vacuum of space and then extinguished will expose unoxidized pyrophoric particles, which can be used to generate sufficient heat to relight the propellant when oxidizer is flowed. The experiments outlined in this work aim to develop a suitable pyrophoric material for use in an RFG, demonstrate pyrophoric relight, and characterize performance under conditions relevant to a hybrid rocket thruster. Magnesium, lithium, calcium, and an alloy of titanium, chromium, and manganese (TiCrMn) were investigated to determine suitability of pure metals as RFG additives. Additionally, aluminum hydride (AlH3), lithium aluminum hydride (LiAlH4), lithium borohydride (LiBH4), and magnesium hydride (MgH2) were investigated to determine suitability of metals hydrides as RFG additives or as precursors for pure-metal RFG additives. Pyrophoric metals have been previously investigated as additives for increasing the regression rate of hybrid fuels, but to the author's knowledge, these materials have not been specifically investigated for their ability to ignite a propellant pyrophorically. Commercial research-grade metals were obtained as coarse powders, then ball-milled to attempt to reduce particle size below a critical diameter needed for pyrophoricity. Magnesium hydride was ball-milled and then cycled in a hydride cycling

  2. Single-element coaxial injector for rocket fuel

    NASA Technical Reports Server (NTRS)

    Larson, L. L.

    1969-01-01

    Improved injector for oxygen difluoride and diborane has better mixing characteristics and is able to project fuel onto the wall of the combustion chamber for better cooling. It produces an essentially conical, diverging, continuous sheet of propellant mixture formed by similarly shaped and continuously impinging sheets of fuel and oxidant.

  3. Fundamental phenomena on fuel decomposition and boundary layer combustion processes with applications to hybrid rocket motors

    NASA Technical Reports Server (NTRS)

    Kuo, Kenneth K.; Lu, Y. C.; Chiaverini, Martin J.; Harting, George C.

    1994-01-01

    An experimental study on the fundamental processes involved in fuel decomposition and boundary layer combustion in hybrid rocket motors is being conducted at the High Pressure Combustion Laboratory of the Pennsylvania State University. This research should provide a useful engineering technology base in the development of hybrid rocket motors as well as a fundamental understanding of the complex processes involved in hybrid propulsion. A high pressure slab motor has been designed and manufactured for conducting experimental investigations. Oxidizer (LOX or GOX) supply and control systems have been designed and partly constructed for the head-end injection into the test chamber. Experiments using HTPB fuel, as well as fuels supplied by NASA designated industrial companies will be conducted. Design and construction of fuel casting molds and sample holders have been completed. The portion of these items for industrial company fuel casting will be sent to the McDonnell Douglas Aerospace Corporation in the near future. The study focuses on the following areas: observation of solid fuel burning processes with LOX or GOX, measurement and correlation of solid fuel regression rate with operating conditions, measurement of flame temperature and radical species concentrations, determination of the solid fuel subsurface temperature profile, and utilization of experimental data for validation of a companion theoretical study (Part 2) also being conducted at PSU.

  4. Fuel Regression Characteristics of Cascaded Multistage Impinging-Jet (CAMUI) Type Hybrid Rocket

    NASA Astrophysics Data System (ADS)

    Itoh, Mitsunori; Maeda, Takenori; Kakikura, Akihito; Kaneko, Yudai; Mori, Kazuhiro; Nakashima, Takuji; Wakita, Masashi; Uematsu, Tsutomu; Totani, Tsuyoshi; Oshima, Nobuyuki; Nagata, Harunori

    A series of lab-scale firing tests was conducted to investigate the fuel regression characteristics of Cascaded Multistage Impinging-jet (CAMUI) type hybrid rocket. The alternative fuel grain used in this rocket consists of a number of cylindrical fuel blocks with two ports, which were aligned along the axis of the combustion chamber with a small gap. The ports are aligned staggered with respect to ones of neighboring blocks so that the combustion gas flow impinges on the forward-end surface of each block. In this fuel grain, forward-end surfaces, back-end surfaces and ports of fuel blocks contribute as burning surfaces. Polyethylene and LOX were used as a propellant, and the tests were conducted at the chamber pressure of 0.5 2MPa and the mass flux of 50 200kg/m2s. Main results obtained in this study are in the followings: The regression rate of each surface was obtained as a function of the propellant mass flux and local equivalent ratio of the combustion gas. At back-end surfaces the regression rate has a high sensitivity on the gap height of neighboring fuel blocks. These fuel regression characteristics will contribute as fundamental data to improve the optimum design of the fuel grain.

  5. Fundamental phenomena on fuel decomposition and boundary layer combustion processes with applications to hybrid rocket motors

    NASA Astrophysics Data System (ADS)

    Kuo, Kenneth K.; Lu, Y. C.; Chiaverini, Martin J.; Harting, George C.

    1994-11-01

    An experimental study on the fundamental processes involved in fuel decomposition and boundary layer combustion in hybrid rocket motors is being conducted at the High Pressure Combustion Laboratory of the Pennsylvania State University. This research should provide a useful engineering technology base in the development of hybrid rocket motors as well as a fundamental understanding of the complex processes involved in hybrid propulsion. A high pressure slab motor has been designed and manufactured for conducting experimental investigations. Oxidizer (LOX or GOX) supply and control systems have been designed and partly constructed for the head-end injection into the test chamber. Experiments using HTPB fuel, as well as fuels supplied by NASA designated industrial companies will be conducted. Design and construction of fuel casting molds and sample holders have been completed. The portion of these items for industrial company fuel casting will be sent to the McDonnell Douglas Aerospace Corporation in the near future. The study focuses on the following areas: observation of solid fuel burning processes with LOX or GOX, measurement and correlation of solid fuel regression rate with operating conditions, measurement of flame temperature and radical species concentrations, determination of the solid fuel subsurface temperature profile, and utilization of experimental data for validation of a companion theoretical study (Part 2) also being conducted at PSU.

  6. Reduction of gaseous pollutant emissions from gas turbine combustors using hydrogen-enriched jet fuel

    NASA Technical Reports Server (NTRS)

    Clayton, R. M.

    1976-01-01

    Recent progress in an evaluation of the applicability of the hydrogen enrichment concept to achieve ultralow gaseous pollutant emission from gas turbine combustion systems is described. The target emission indexes for the program are 1.0 for oxides of nitrogen and carbon monoxide, and 0.5 for unburned hydrocarbons. The basic concept utilizes premixed molecular hydrogen, conventional jet fuel, and air to depress the lean flammability limit of the mixed fuel. This is shown to permit very lean combustion with its low NOx production while simulataneously providing an increased flame stability margin with which to maintain low CO and HC emission. Experimental emission characteristics and selected analytical results are presented for a cylindrical research combustor designed for operation with inlet-air state conditions typical for a 30:1 compression ratio, high bypass ratio, turbofan commercial engine.

  7. Liquefied Gaseous Fuels Safety and Environmental Control Assessment Program: second status report

    SciTech Connect

    Not Available

    1980-10-01

    The Assistant Secretary for Environment has responsibility for identifying, characterizing, and ameliorating the environmental, health, and safety issues and public concerns associated with commercial operation of specific energy systems. The need for developing a safety and environmental control assessment for liquefied gaseous fuels was identified by the Environmental and Safety Engineering Division as a result of discussions with various governmental, industry, and academic persons having expertise with respect to the particular materials involved: liquefied natural gas, liquefied petroleum gas, hydrogen, and anhydrous ammonia. This document is arranged in three volumes and reports on progress in the Liquefied Gaseous Fuels (LGF) Safety and Environmental Control Assessment Program made in Fiscal Year (FY)-1979 and early FY-1980. Volume 1 (Executive Summary) describes the background, purpose and organization of the LGF Program and contains summaries of the 25 reports presented in Volumes 2 and 3. Annotated bibliographies on Liquefied Natural Gas (LNG) Safety and Environmental Control Research and on Fire Safety and Hazards of Liquefied Petroleum Gas (LPG) are included in Volume 1.

  8. Experimental research on the rotating detonation in gaseous fuels-oxygen mixtures

    NASA Astrophysics Data System (ADS)

    Kindracki, J.; Wolański, P.; Gut, Z.

    2011-04-01

    An experimental study on rotating detonation is presented in this paper. The study was focused on the possibility of using rotating detonation in a rocket engine. The research was divided into two parts: the first part was devoted to obtaining the initiation of rotating detonation in fuel-oxygen mixture; the second was aimed at determination of the range of propagation stability as a function of chamber pressure, composition, and geometry. Additionally, thrust and specific impulse were determined in the latter stage. In the paper, only rich mixture is described, because using such a composition in rocket combustion chambers maximizes the specific impulse and thrust. In the experiments, two kinds of geometry were examined: cylindrical and cylindrical-conic, the latter can be simulated by a simple aerospike nozzle. Methane, ethane, and propane were used as fuel. The pressure-time courses in the manifolds and in the chamber are presented. The thrust-time profile and detonation velocity calculated from measured pressure peaks are shown. To confirm the performance of a rocket engine with rotating detonation as a high energy gas generator, a model of a simple engine was designed, built, and tested. In the tests, the model of the engine was connected to the dump tank. This solution enables different environmental conditions from a range of flight from 16 km altitude to sea level to be simulated. The obtained specific impulse for pressure in the chamber of max. 1.2 bar and a small nozzle expansion ratio of about 3.5 was close to 1,500 m/s.

  9. Atmospheric Dispersion of Hypergolic Liquid Rocket Fuels. Volume 1

    DTIC Science & Technology

    1984-11-01

    3) 2. Excess Oxidant Reactions ..... .............. .. 37 C. Calctilation of Fireball Size and Quantification of Heat Flux...this hypergolic fuel-oxidizer combustion have been reported in the literature. While the identification and quantification of all such chemical...due to thermal or toxicity considerations. Dimethylnitrosamine (NDMA), for example, is an expected and confirmed product from the nitrogen tetroxide

  10. High-pressure soot formation and diffusion flame extinction characteristics of gaseous and liquid fuels

    NASA Astrophysics Data System (ADS)

    Karatas, Ahmet Emre

    High-pressure soot formation and flame stability characteristics were studied experimentally in laminar diffusion flames. For the former, radially resolved soot volume fraction and temperature profiles were measured in axisymmetric co-flow laminar diffusion flames of pre-vaporized n-heptane-air, undiluted ethylene-air, and nitrogen and carbon dioxide diluted ethylene-air at elevated pressures. Abel inversion was used to re-construct radially resolved data from the line-of-sight spectral soot emission measurements. For the latter, flame extinction strain rate was measured in counterflow laminar diffusion flames of C1-4 alcohols and hydrocarbon fuels of n-heptane, n-octane, iso-octane, toluene, Jet-A, and biodiesel. The luminous flame height, as marked by visible soot radiation, of the nitrogen- and helium-diluted n-heptane and nitrogen- and carbon dioxide-diluted ethylene flames stayed constant at all pressures. In pure ethylene flames, flame heights initially increased with pressure, but changed little above 5 atm. The maximum soot yield as a function of pressure in nitrogen-diluted n-heptane diffusion flames indicate that n-heptane flames are slightly more sensitive to pressure than gaseous alkane hydrocarbon flames at least up to 7 atm. Ethylene's maximum soot volume fractions were much higher than those of ethane and n-heptane diluted with nitrogen (fuel to nitrogen mass flow ratio is about 0.5). Pressure dependence of the peak carbon conversion to soot, defined as the percentage of fuel's carbon content converted to soot, was assessed and compared to previous measurements with other gaseous fuels. Maximum soot volume fractions were consistently lower in carbon dioxide-diluted flames between 5 and 15 atm but approached similar values to those in nitrogen-diluted flames at 20 atm. This observation implies that the chemical soot suppression effect of carbon dioxide, previously demonstrated at atmospheric pressure, is also present at elevated pressures up to 15 atm

  11. Metallized Gelled Propellants: Heat Transfer of a Rocket Engine Fueled by Oxygen/RP-1/Aluminum Was Measured by a Calorimeter

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan A.

    1998-01-01

    A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. These analyses used data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber, and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt % loadings of aluminum (Al) with gaseous oxygen as the oxidizer. Heat transfer measurements were made with a calorimeter chamber and nozzle setup that had a total of 31 cooling channels. A gelled fuel coating, composed of unburned gelled fuel and partially combusted RP-1, formed in the 0-, 5- and 55-wt % engines. For the 0- and 5-wt % RP-1/Al, the coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber. This heat flux reduction was analyzed by comparing engine firings and the changes in the heat flux during a firing at NASA Lewis Research Center's Rocket Laboratories. This work is part of an ongoing series of analyses of metallized gelled propellants.

  12. Destruction of explosives and rocket fuels by supercritical water oxidation

    SciTech Connect

    Dyer, R.B.; Buelow, S.J.; Harradine, D.M.; Robinson, J.M.; Foy, B.R.; Atencio, J.H.; Dell'Orco, P.C.; Funk, K.A.; McInroy, R.E.; Rofer, C.K.; Counce, D.A.; Trujillo, P.E. Jr. ); Wander, J.D. )

    1992-01-01

    Traditional methods for disposing of PEPs have been open burning or open detonation (OB/OD); however, regulatory agencies are likely to prohibit OB/OD because of the uncontrolled air emissions and soil contaminations. Likewise, controlled incineration carries a liability for air pollution because large quantities of NO{sub x} are produced in the conventional combustion chemistry of PEPS. Soil and ground water have already been contaminated with PEPs through normal operations at manufacturing plants and military bases. Incineration can be used for decontamination of these soils, with the associated liability for air pollution, but few satisfactory and economic methods exist for ground water decontamination. A clear need exists for improved disposal and destruction methods. The destruction of energetic materials, including propellants, explosives and pyrotechnics (PEPS) by oxidation in supercritical water is described. The focus is on the chemistry of the process. The destruction efficiencies and products of reaction contained in the aqueous and gaseous effluents of several representative PEPs are reported.

  13. Self-oscillations of an unstable fuel combustion in the combustion chamber of a liquid-propellant rocket engine

    NASA Astrophysics Data System (ADS)

    Gotsulenko, V. V.; Gotsulenko, V. N.

    2013-01-01

    The form of the self-oscillations of a vibrating combustion of a fuel in the combustion chamber of a liquidpropellant rocket engine, caused by the fuel-combustion lag and the heat release, was determined. The character of change in these self-oscillations with increase in the time of the fuel-combustion lag was investigated.

  14. Fundamental phenomena on fuel decomposition and boundary layer combustion processes with applications to hybrid rocket motors

    NASA Technical Reports Server (NTRS)

    Kuo, Kenneth K.; Lu, Y. C.; Chiaverini, Martin J.; Harting, George C.

    1994-01-01

    An experimental study on the fundamental processes involved in fuel decomposition and boundary layer combustion in hybrid rocket motors is being conducted at the High Pressure Combustion Laboratory of the Pennsylvania State University. This research should provide an engineering technology base for development of large scale hybrid rocket motors as well as a fundamental understanding of the complex processes involved in hybrid propulsion. A high pressure slab motor has been designed for conducting experimental investigations. Oxidizer (LOX or GOX) is injected through the head-end over a solid fuel (HTPB) surface. Experiments using fuels supplied by NASA designated industrial companies will also be conducted. The study focuses on the following areas: measurement and observation of solid fuel burning with LOX or GOX, correlation of solid fuel regression rate with operating conditions, measurement of flame temperature and radical species concentrations, determination of the solid fuel subsurface temperature profile, and utilization of experimental data for validation of a companion theoretical study also being conducted at PSU.

  15. Computing Q-D Relationships for Storage of Rocket Fuels

    NASA Technical Reports Server (NTRS)

    Jester, Keith

    2005-01-01

    The Quantity Distance Measurement Tool is a GIS BASEP computer program that aids safety engineers by calculating quantity-distance (Q-D) relationships for vessels that contain explosive chemicals used in testing rocket engines. (Q-D relationships are standard relationships between specified quantities of specified explosive materials and minimum distances by which they must be separated from persons, objects, and other explosives to obtain specified types and degrees of protection.) The program uses customized geographic-information-system (GIS) software and calculates Q-D relationships in accordance with NASA's Safety Standard For Explosives, Propellants, and Pyrotechnics. Displays generated by the program enable the identification of hazards, showing the relationships of propellant-storage-vessel safety buffers to inhabited facilities and public roads. Current Q-D information is calculated and maintained in graphical form for all vessels that contain propellants or other chemicals, the explosiveness of which is expressed in TNT equivalents [amounts of trinitrotoluene (TNT) having equivalent explosive effects]. The program is useful in the acquisition, siting, construction, and/or modification of storage vessels and other facilities in the development of an improved test-facility safety program.

  16. Aluminum-fueled rockets for the space transportation system

    NASA Technical Reports Server (NTRS)

    Cutler, Andrew Hall

    1992-01-01

    Aluminum-fueled engines, used to propel orbital transfer vehicles (OTV's), offer benefits to the Space Transportation System (STS) if scrap aluminum can be scavenged at a reasonable cost. Aluminum scavenged from Space Shuttle external tanks could replace propellants hauled from Earth, thus allowing more payloads to be sent to their final destinations at the same Shuttle launch rate. To allow OTV use of aluminum fuel, two new items would be required: a facility to reprocess aluminum from external tanks and an engine for the OTV which could burn aluminum. Design of the orbital transfer vehicle would have to differ substantially from current concepts for it to carry and use the aluminum fuel. The aluminum reprocessing facility would probably have a mass of under 15 metric tons and would probably cost less that $200,000,000. Development of an aluminum-burning engine would no doubt be extremely expensive (1 to 2 billion dollars), but this amount would be adequately repaid by increased STS throughput. Engine production cost is difficult to estimate, but even an extremely high cost (e.g., $250,000,000 per engine) would not significantly increase orbit-raising expenses.

  17. Soil factors of ecosystems' disturbance risk reduction under the impact of rocket fuel

    NASA Astrophysics Data System (ADS)

    Krechetov, Pavel; Koroleva, Tatyana; Sharapova, Anna; Chernitsova, Olga

    2016-04-01

    Environmental impacts occur at all stages of space rocket launch. One of the most dangerous consequences of a missile launch is pollution by components of rocket fuels ((unsymmetrical dimethylhydrazine (UDMH)). The areas subjected to falls of the used stages of carrier rockets launched from the Baikonur cosmodrome occupy thousands of square kilometers of different natural landscapes: from dry steppes of Kazakhstan to the taiga of West Siberia and mountains of the Altai-Sayany region. The study aims at assessing the environmental risk of adverse effects of rocket fuel on the soil. Experimental studies have been performed on soil and rock samples with specified parameters of the material composition. The effect of organic matter, acid-base properties, particle size distribution, and mineralogy on the decrease in the concentration of UDMH in equilibrium solutions has been studied. It has been found that the soil factors are arranged in the following series according to the effect on UDMH mobility: acid-base properties > organic matter content >clay fraction mineralogy > particle size distribution. The estimation of the rate of self-purification of contaminated soil is carried out. Experimental study of the behavior of UDMH in soil allowed to define a model for calculating critical loads of UDMH in terrestrial ecosystems.

  18. Formulation and Testing of Paraffin-Based Solid Fuels Containing Energetic Additives for Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Larson, Daniel B.; Boyer, Eric; Wachs,Trevor; Kuo, Kenneth K.; Story, George

    2012-01-01

    Many approaches have been considered in an effort to improve the regression rate of solid fuels for hybrid rocket applications. One promising method is to use a fuel with a fast burning rate such as paraffin wax; however, additional performance increases to the fuel regression rate are necessary to make the fuel a viable candidate to replace current launch propulsion systems. The addition of energetic and/or nano-sized particles is one way to increase mass-burning rates of the solid fuels and increase the overall performance of the hybrid rocket motor.1,2 Several paraffin-based fuel grains with various energetic additives (e.g., lithium aluminum hydride (LiAlH4) have been cast in an attempt to improve regression rates. There are two major advantages to introducing LiAlH4 additive into the solid fuel matrix: 1) the increased characteristic velocity, 2) decreased dependency of Isp on oxidizer-to-fuel ratio. The testing and characterization of these solid-fuel grains have shown that continued work is necessary to eliminate unburned/unreacted fuel in downstream sections of the test apparatus.3 Changes to the fuel matrix include higher melting point wax and smaller energetic additive particles. The reduction in particle size through various methods can result in more homogeneous grain structure. The higher melting point wax can serve to reduce the melt-layer thickness, allowing the LiAlH4 particles to react closer to the burning surface, thus increasing the heat feedback rate and fuel regression rate. In addition to the formulation of LiAlH4 and paraffin wax solid-fuel grains, liquid additives of triethylaluminum and diisobutylaluminum hydride will be included in this study. Another promising fuel formulation consideration is to incorporate a small percentage of RDX as an additive to paraffin. A novel casting technique will be used by dissolving RDX in a solvent to crystallize the energetic additive. After dissolving the RDX in a solvent chosen for its compatibility

  19. System Modeling and Diagnostics for Liquefying-Fuel Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Poll, Scott; Iverson, David; Ou, Jeremy; Sanderfer, Dwight; Patterson-Hine, Ann

    2003-01-01

    A Hybrid Combustion Facility (HCF) was recently built at NASA Ames Research Center to study the combustion properties of a new fuel formulation that burns approximately three times faster than conventional hybrid fuels. Researchers at Ames working in the area of Integrated Vehicle Health Management recognized a good opportunity to apply IVHM techniques to a candidate technology for next generation launch systems. Five tools were selected to examine various IVHM techniques for the HCF. Three of the tools, TEAMS (Testability Engineering and Maintenance System), L2 (Livingstone2), and RODON, are model-based reasoning (or diagnostic) systems. Two other tools in this study, ICS (Interval Constraint Simulator) and IMS (Inductive Monitoring System) do not attempt to isolate the cause of the failure but may be used for fault detection. Models of varying scope and completeness were created, both qualitative and quantitative. In each of the models, the structure and behavior of the physical system are captured. In the qualitative models, the temporal aspects of the system behavior and the abstraction of sensor data are handled outside of the model and require the development of additional code. In the quantitative model, less extensive processing code is also necessary. Examples of fault diagnoses are given.

  20. Combustion performance of bipropellant liquid rocket engine combustors with fuel-impingement cooling

    NASA Astrophysics Data System (ADS)

    Jiang, Tsung Leo; Chiang, Wei-Tang; Jang, Shyh-Dihng

    1995-05-01

    In order to obtain an accurate combustion analyses which are important in the thruster design of modern advanced liquid rocket engine, flow analysis should be conducted from the injector phase down to the propulsive nozzle throat. Thus, in the present study, flow analysis for the axisymmetric thrust chamber of an OMV(exp 3) installed with a pintle-type ring-shaped injector and a conical convergent nozzle is conducted. Liquid monomethyl hydrazine (MMH) and nitrogen tetroxide (NTO) storable bipropellants are used as fuel and oxidizer sources. An optimum injected fuel and oxidizer droplet-size combination is proposed. Finally, the results obtained are presented.

  1. Combustion performance of bipropellant liquid rocket engine combustors with fuel-impingement cooling

    SciTech Connect

    Jiang, T.L.; Chiang, W.; Jang, S.

    1995-05-01

    In order to obtain an accurate combustion analyses which are important in the thruster design of modern advanced liquid rocket engine, flow analysis should be conducted from the injector phase down to the propulsive nozzle throat. Thus, in the present study, flow analysis for the axisymmetric thrust chamber of an OMV(exp 3) installed with a pintle-type ring-shaped injector and a conical convergent nozzle is conducted. Liquid monomethyl hydrazine (MMH) and nitrogen tetroxide (NTO) storable bipropellants are used as fuel and oxidizer sources. An optimum injected fuel and oxidizer droplet-size combination is proposed. Finally, the results obtained are presented. 4 refs.

  2. Liquefied gaseous fuels safety and environmental control assessment program: third status report

    SciTech Connect

    Not Available

    1982-03-01

    This Status Report contains contributions from all contractors currently participating in the DOE Liquefied Gaseous Fuels (LG) Safety and Environmental Control Assessment Program and is presented in two principal sections. Section I is an Executive Summary of work done by all program participants. Section II is a presentation of fourteen individual reports (A through N) on specific LGF Program activities. The emphasis of Section II is on research conducted by Lawrence Livermore National Laboratory (Reports A through M). Report N, an annotated bibliography of literature related to LNG safety and environmental control, was prepared by Pacific Northwest Laboratory (PNL) as part of its LGF Safety Studies Project. Other organizations who contributed to this Status Report are Aerojet Energy Conversion Company; Applied Technology Corporation; Arthur D. Little, Incorporated; C/sub v/ International, Incorporated; Institute of Gas Technology; and Massachusetts Institute of Technology. Separate abstracts have been prepared for Reports A through N for inclusion in the Energy Data Base.

  3. Intermediate pyrolysis of biomass energy pellets for producing sustainable liquid, gaseous and solid fuels.

    PubMed

    Yang, Y; Brammer, J G; Mahmood, A S N; Hornung, A

    2014-10-01

    This work describes the use of intermediate pyrolysis system to produce liquid, gaseous and solid fuels from pelletised wood and barley straw feedstock. Experiments were conducted in a pilot-scale system and all products were collected and analysed. The liquid products were separated into an aqueous phase and an organic phase (pyrolysis oil) under gravity. The oil yields were 34.1 wt.% and 12.0 wt.% for wood and barley straw, respectively. Analysis found that both oils were rich in heterocyclic and phenolic compounds and have heating values over 24 MJ/kg. The yields of char for both feedstocks were found to be about 30 wt.%, with heating values similar to that of typical sub-bituminous class coal. Gas yields were calculated to be approximately 20 wt.%. Studies showed that both gases had heating values similar to that of downdraft gasification producer gas. Analysis on product energy yields indicated the process efficiency was about 75%.

  4. Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter

    NASA Technical Reports Server (NTRS)

    Messing, Wesley E.

    1959-01-01

    Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

  5. A preliminary assessment of the feasibility of deriving liquid and gaseous fuels from grown and waste organics

    NASA Technical Reports Server (NTRS)

    Graham, R. W.; Reynolds, T. W.; Hsu, Y. Y.

    1976-01-01

    The anticipated depletion of our resources of natural gas and petroleum in a few decades has caused a search for renewable sources of fuel. Among the possibilities is the chemical conversion of waste and grown organic matter into gaseous or liquid fuels. The overall feasibility of such a system is considered from the technical, economic, and social viewpoints. Although there are a number of difficult problems to overcome, this preliminary study indicates that this option could provide between 4 and 10 percent of the U.S. energy needs. Estimated costs of fuels derived from grown organic material are appreciably higher than today's market price for fossil fuel. The cost of fuel derived from waste organics is competitive with fossil fuel prices. Economic and social reasons will prohibit the allocation of good food producing land to fuel crop production.

  6. A Versatile Rocket Engine Hot Gas Facility

    NASA Technical Reports Server (NTRS)

    Green, James M.

    1993-01-01

    The capabilities of a versatile rocket engine facility, located in the Rocket Laboratory at the NASA Lewis Research Center, are presented. The gaseous hydrogen/oxygen facility can be used for thermal shock and hot gas testing of materials and structures as well as rocket propulsion testing. Testing over a wide range of operating conditions in both fuel and oxygen rich regimes can be conducted, with cooled or uncooled test specimens. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods with rapid turnaround between programs.

  7. Experimental investigation of paraffin-based fuels for hybrid rocket propulsion

    NASA Astrophysics Data System (ADS)

    Galfetti, L.; Merotto, L.; Boiocchi, M.; Maggi, F.; DeLuca, L. T.

    2013-03-01

    Solid fuels for hybrid rockets were characterized in the framework of a research project aimed to develop a new generation of solid fuels, combining at the same time good mechanical and ballistic properties. Original techniques were implemented in order to improve paraffin-based fuels. The first strengthening technique involves the use of a polyurethane foam (PUF); a second technique is based on thermoplastic polymers mixed at molecular level with the paraffin binder. A ballistic characterization of paraffin-based hybrid rocket solid fuels was performed, considering pure wax-based fuels and fuels doped with suitable metal additives. Nano-Al powders and metal hydrides (magnesium hydride (MgH2), lithium aluminum hydride (LiAlH4 )) were used as fillers in paraffin matrices. The results of this investigation show a strong correlation between the measured viscosity of the melted paraffin layer and the regression rate: a decrease of viscosity increases the regression rate. This trend is due to the increasing development of entrainment phenomena, which strongly increase the regression rate. Addition of LiAlH4 (mass fraction 10%) can further increase the regression rate up to 378% with respect to the pure HTPB regression rate, taken as baseline reference fuel. The highest regression rates were found for the Solid Wax (SW) composition, added with 5% MgH2 mass fraction; at 350 kg/(m2s) oxygen mass flux, the measured regression rate, averaged in space and time, was 2.5 mm/s, which is approximately five times higher than that of the pure HTPB composition. Compositions added with nanosized aluminum powders were compared with those added with MgH2, using gel or solid wax.

  8. Reacting shock waves characteristics for biogas compared to other gaseous fuel

    NASA Astrophysics Data System (ADS)

    Wahid, Mazlan Abdul; Ujir, Haffis

    2012-06-01

    Present article aims to report an experimental study conducted to characterize the reacting shock waves for biogas compared to several other gaseous fuels. A dedicated experimental system which consists of a stainless steel tube with inner diameter of 100mm, a data acquisition system, ignition control unit and gas filling system was built in order to measure the characteristics of high speed reacting shock waves for synthetic biogas such as, pressure history, velocity and cell width. Two types of hydrocarbon fuels were used for comparison in this investigation; propane and natural gas with 92.7% methane. Biogas was synthetically produced by mixing 65% natural gas with 35% carbon dioxide. The oxygen concentration in the oxidizer mixture was diluted with nitrogen gas at various percentage of dilution. Results show that natural gas and biogas were not sensitive to detonation propagation compared to propane. For biogas, methane, and propane it was found that in smooth inner-wall tube, detonation will likely to occur if the percent of dilution gas is not more than approximately 8%, 10% and 35%, respectively. In order to decrease the tube length required for deflagration to detonation transition, an array of obstacles with identical blockage ratio was placed inside the tube near the ignition source. The effect of combustion wave-obstacle interaction was also investigated.

  9. Experimental investigation of anodic gaseous concentration of a practical seal-less solid oxide fuel cell

    NASA Astrophysics Data System (ADS)

    Momma, Akihiko; Kaga, Yasuo; Takano, Kiyonami; Nozaki, Ken; Negishi, Akira; Kato, Ken; Kato, Tohru; Inagaki, Toru; Yoshida, Hiroyuki; Hoshino, Koji; Yamada, Masaharu; Akbay, Taner; Akikusa, Jun

    In order to verify the validity of the simulation and to investigate the gaseous diffusion from the outlet of the anode, anodic gas concentration measurements of a seal-less disk-type solid oxide fuel cell (SOFC) were carried out using quadrupole mass spectrometer (QMS). Simultaneous gas sampling was conducted from the five sampling ports made at the anode separator. The uniformity of the radial gas flow in the anode was confirmed by analyzing the gas from four sampling ports located at a concentric circle. H 2, H 2O and N 2 concentration profiles were measured and simulated under various fuel utilization ( Uf) conditions and changing the gas flow rate. The diffusion of N 2 into the anode was found to become less with increasing Uf owing to the lesser diffusivity of N 2 in H 2O than in H 2. From the simulation, the existence of the reverse current, i.e., electrolysis current, in the outlet region was predicted. It was confirmed that the existence of the electrolysis current is possible by measuring the concentration of the gas in the anode under electrolysis operations. The comparison of V- i characteristics measured and simulated revealed that the effect of the concentration polarization is not significant and suggested the validity of the assumption made for the simulation.

  10. Turbulent mixing of multiple co-axial gaseous fuel jets in a supersonic airstream

    NASA Technical Reports Server (NTRS)

    Lorber, A. K.; Schetz, J. A.

    1974-01-01

    An experimental and analytical investigation of a strut-mounted, four-nozzle, downstream-facing, gaseous fuel-injector assembly was conducted in a Mach 4 airstream at 154 psia and 520 R. Helium was used as the injectant in order to simulate a low molecular weight fuel, and the interjet spacing (S/D) was the main parameter varied. The principal data are in the form of helium concentration profiles at seven axial stations, Mach number distributions at three axial stations, and Schlieren photographs of the flow field at different interjet spacings. An approximate analysis was developed based upon a linearization in the Von Mises plane, Crocco integrals for the temperature and concentrations fields, and an eddy viscosity model. Good agreement with the data was achieved. It was found that, at these conditions, interjet spacing has a significant effect on mixing only up to S/D of approximately 2.5, with the concentration of the injectant at a given point becoming smaller as S/D is increased.

  11. Turbulent mixing of multiple co-axial gaseous fuel jets in a supersonic airstream

    NASA Technical Reports Server (NTRS)

    Lorber, A. K.; Schetz, J. A.

    1974-01-01

    An experimental and analytical investigation of a strut-mounted, four-nozzle, downstream-facing, gaseous fuel-injector assembly was conducted in a Mach 4 airstream at 154 psia and 520 R. Helium was used as the injectant in order to simulate a low molecular weight fuel, and the interjet spacing (S/D) was the main parameter varied. The principal data are in the form of helium concentration profiles at seven axial stations, Mach number distributions at three axial stations, and Schlieren photographs of the flow field at different interjet spacings. An approximate analysis was developed based upon a linearization in the Von Mises plane, Crocco integrals for the temperature and concentrations fields, and an eddy viscosity model. Good agreement with the data was achieved. It was found that, at these conditions, interjet spacing has a significant effect on mixing only up to S/D of approximately 2.5, with the concentration of the injectant at a given point becoming smaller as S/D is increased.

  12. Numerical investigation on the regression rate of hybrid rocket motor with star swirl fuel grain

    NASA Astrophysics Data System (ADS)

    Zhang, Shuai; Hu, Fan; Zhang, Weihua

    2016-10-01

    Although hybrid rocket motor is prospected to have distinct advantages over liquid and solid rocket motor, low regression rate and insufficient efficiency are two major disadvantages which have prevented it from being commercially viable. In recent years, complex fuel grain configurations are attractive in overcoming the disadvantages with the help of Rapid Prototyping technology. In this work, an attempt has been made to numerically investigate the flow field characteristics and local regression rate distribution inside the hybrid rocket motor with complex star swirl grain. A propellant combination with GOX and HTPB has been chosen. The numerical model is established based on the three dimensional Navier-Stokes equations with turbulence, combustion, and coupled gas/solid phase formulations. The calculated fuel regression rate is compared with the experimental data to validate the accuracy of numerical model. The results indicate that, comparing the star swirl grain with the tube grain under the conditions of the same port area and the same grain length, the burning surface area rises about 200%, the spatially averaged regression rate rises as high as about 60%, and the oxidizer can combust sufficiently due to the big vortex around the axis in the aft-mixing chamber. The combustion efficiency of star swirl grain is better and more stable than that of tube grain.

  13. Design Considerations of ISTAR Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

    NASA Technical Reports Server (NTRS)

    Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

    2003-01-01

    The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system that produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

  14. Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

    NASA Technical Reports Server (NTRS)

    Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

    2002-01-01

    The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

  15. Gaseous fission product management for molten salt reactors and vented fuel systems

    SciTech Connect

    Messenger, S. J.; Forsberg, C.; Massie, M.

    2012-07-01

    Fission gas disposal is one of the unresolved difficulties for Molten Salt Reactors (MSRs) and advanced reactors with vented fuel systems. As these systems operate, they produce many radioactive isotopes of xenon and krypton (e.g. {sup 135}Xe t{sub 1/2} = 9.14 hours and {sup 85}Kr t{sub 1/2}= 10.73 years). Removing these gases proves vital to the success of such reactor designs for two reasons. First, the gases act as large neutron sinks which decrease reactivity and must be counterbalanced by increasing fuel loading. Second, for MSRs, inert fission product gases naturally separate quickly from high temperature salts, thus creating high vapor pressure which poses safety concerns. For advanced reactors with solid vented fuel, the gases are allowed to escape into an off-gas system and thus must be managed. Because of time delays in transport of fission product gases in vented fuel systems, some of the shorter-lived radionuclides will decay away thereby reducing the fission gas source term relative to an MSR. To calculate the fission gas source term of a typical molten salt reactor, we modeled a 1000 MWe graphite moderated thorium MSR similar to that detailed in Mathieu et al. [1]. The fuel salt used in these calculations was LiF (78 mole percent) - (HN)F 4 (22 mole percent) with a heavy nuclide composition of 3.86% {sup 233}U and 96.14% {sup 232}Th by mass. Before we can remove the fission product gases produced by this reactor configuration, we must first develop an appropriate storage mechanism. The gases could be stored in pressurized containers but then one must be concerned about bottle failure. Methods to trap noble gases in matrices are expensive and complex. Alternatively, there are direct storage/disposal options: direct injection into the Earth or injecting a grout-based product into the Earth. Advances in drilling technologies, hydro fracture technologies, and methods for the sequestration of carbon dioxide from fossil fuel plants are creating new options

  16. Low-temperature Ignition-delay Characteristics of Several Rocket Fuels with Mixed Acid in Modified Open-cup-type Apparatus

    NASA Technical Reports Server (NTRS)

    Miller, Riley O

    1950-01-01

    Summaries of low-temperature self-ignition data of various rocket fuels with mixed acid (nitric plus sulfuric) are presented. Several fuels are shown to have shorter ignition-delay intervals and less variation in delay intervals at moderate and sub-zero temperatures than crude N-ethylaniline (monoethylaniline),a rocket fuel in current use.

  17. Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

    NASA Technical Reports Server (NTRS)

    Pavli, A. J.

    1979-01-01

    The feasibility of using a heavy hydrocarbon fuel as a rocket propellant is examined. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. Experiments were done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in) to 55.9 cm (22 in). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by reaming each injector several times to provide test data over a range of injector pressure drop.

  18. Towards Safer Rocket Fuels: Hypergolic Imidazolylidene-Borane Compounds as Replacements for Hydrazine Derivatives.

    PubMed

    Huang, Shi; Qi, Xiujuan; Liu, Tianlin; Wang, Kangcai; Zhang, Wenquan; Li, Jianlin; Zhang, Qinghua

    2016-07-11

    Currently, toxic and volatile hydrazine derivatives are still the main fuel choices for liquid bipropellants, especially in some traditional rocket propulsion systems. Therefore, the search for safer hypergolic fuels as replacements for hydrazine derivatives has been one of the most challenging tasks. In this study, six imidazolylidene-borane compounds with zwitterionic structure have been synthesized and characterized, and their hypergolic reactivity has been studied. As expected, these compounds exhibited fast spontaneous combustion upon contact with white fuming nitric acid (WFNA). Among them, compound 5 showed excellent integrated properties including wide liquid operating range (-70-160 °C), superior loading density (0.99 g cm(-3) ), ultrafast ignition delay times with WFNA (15 ms), and high specific impulse (303.5 s), suggesting promising application potential as safer hypergolic fuels in liquid bipropellant formulations. © 2016 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

  19. Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

    NASA Technical Reports Server (NTRS)

    Pavli, A. J.

    1979-01-01

    An experimental program to determine the feasibility of using a heavy hydrocarbon fuel as a rocket propellant is reported herein. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. The work was done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in.) to 55.9 cm (22 in.). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by 'reaming' each injector several times to provide test data over a range of injector pressure drop.

  20. Computational fluid dynamic simulations of chemical looping fuel reactors utilizing gaseous fuels

    SciTech Connect

    Mahalatkar, Kartikeya; Kuhlman, John; Huckaby, E. David; O'Brien, Thomas

    2011-02-01

    A computational fluid dynamic (CFD) model for the fuel reactor of chemical looping combustion technology has been developed, with special focus on accurately representing the heterogeneous chemical reactions. A continuum two-fluid model was used to describe both the gas and solid phases. Detailed sub-models to account for fluid–particle and particle–particle interaction forces were also incorporated. Two experimental cases were analyzed in this study (Son and Kim, 2006; Mattison et al., 2001). Simulations were carried out to test the capability of the CFD model to capture changes in outlet gas concentrations with changes in number of parameters such as superficial velocity, metal oxide concentration, reactor temperature, etc. For the experiments of Mattisson et al. (2001), detailed time varying outlet concentration values were compared, and it was found that CFD simulations provided a reasonable match with this data.

  1. Annual book of ASTM Standards 2005. Section Five. Petroleum products, lubricants, and fossil fuels. Volume 05.06. Gaseous fuels; coal and coke

    SciTech Connect

    2005-09-15

    The first part covers standards for gaseous fuels. The standard part covers standards on coal and coke including the classification of coals, determination of major elements in coal ash and trace elements in coal, metallurgical properties of coal and coke, methods of analysis of coal and coke, petrographic analysis of coal and coke, physical characteristics of coal, quality assurance and sampling.

  2. Annual book of ASTM Standards 2008. Section Five. Petroleum products, lubricants, and fossil fuels. Volume 05.06. Gaseous fuels; coal and coke

    SciTech Connect

    2008-09-15

    The first part covers standards for gaseous fuels. The second part covers standards on coal and coke including the classification of coals, determination of major elements in coal ash and trace elements in coal, metallurgical properties of coal and coke, methods of analysis of coal and coke, petrogrpahic analysis of coal and coke, physical characteristics of coal, quality assurance and sampling.

  3. Early Rockets

    NASA Image and Video Library

    1940-03-21

    Goddard rocket in launching tower at Roswell, New Mexico, March 21, 1940. Fuel was injected by pumps from the fueling platform at left. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets, which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

  4. Small rocket research and technology

    NASA Technical Reports Server (NTRS)

    Schneider, Steven; Biaglow, James

    1993-01-01

    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

  5. Small rocket research and technology

    NASA Astrophysics Data System (ADS)

    Schneider, Steven; Biaglow, James

    1993-11-01

    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

  6. Calculation of Free-Atom Fractions in Hydrocarbon-Fueled Rocket Engine Plume

    NASA Technical Reports Server (NTRS)

    Verma, Satyajit

    2006-01-01

    Free atom fractions (Beta) of nine elements are calculated in the exhaust plume of CH4- oxygen and RP-1-oxygen fueled rocket engines using free energy minimization method. The Chemical Equilibrium and Applications (CEA) computer program developed by the Glenn Research Center, NASA is used for this purpose. Data on variation of Beta in both fuels as a function of temperature (1600 K - 3100 K) and oxygen to fuel ratios (1.75 to 2.25 by weight) is presented in both tabular and graphical forms. Recommendation is made for the Beta value for a tenth element, Palladium. The CEA computer code was also run to compare with experimentally determined Beta values reported in literature for some of these elements. A reasonable agreement, within a factor of three, between the calculated and reported values is observed. Values reported in this work will be used as a first approximation for pilot rocket engine testing studies at the Stennis Space Center for at least six elements Al, Ca, Cr, Cu, Fe and Ni - until experimental values are generated. The current estimates will be improved when more complete thermodynamic data on the remaining four elements Ag, Co, Mn and Pd are added to the database. A critique of the CEA code is also included.

  7. [Microbiological degradation of asymmetrical dimethylhydrazine--a toxic component of rocket fuel].

    PubMed

    Chugunov, V A; Martovetskaia, I I; Mironova, R I; Fomchenkov, V M; Kholodenko, V P

    2000-01-01

    A possibility of microbiological cleaning of water and soil polluted with asymmetric dimethylhydrazine (ADMH), a highly toxic rocket fuel ingredient (RFI), was studied. Several isolates (bacteria, yeast, and micromycetes) capable of utilizing ADMH as the only source of nitrogen, carbon, and energy were isolated from RFI-polluted tundra soil. Acceleration of RFI biodegradation was achieved using a biosorbent that involved cells of the degrader strain immobilized on granulated activated carbon. Biological testing in Escherichia coli and cereals (wheat and barley) demonstrated that biodegradation significantly decreased the integral toxicity of solutions containing ADMH, suggesting its utility for microbiological cleaning of polluted territories.

  8. Test program to provide confidence in liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

    NASA Technical Reports Server (NTRS)

    Armstrong, Elizabeth S.

    1986-01-01

    In previous tests of liquid oxygen cooling of hydrocarbon fueled rocket engines, small oxygen leaks developed at the throat of the thrust chamber and film cooled the hot gas side of the chamber wall without resulting in catastrophic failure. However, more testing is necessary to demonstrate that a catastropic failure would not occur if cracks developed further upstream between the injector and the throat, where the boundary layer has not been established. Since under normal conditions cracks are expected to form in the throat region of the thrust chamber, cracks must be initiated artificially in order to control their location. Several methods of crack initiation are discussed here.

  9. Test program to provide confidence in liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

    NASA Technical Reports Server (NTRS)

    Armstrong, Elizabeth S.

    1986-01-01

    In previous tests of liquid oxygen cooling of hydrocarbon fueled rocket engines, small oxygen leaks developed at the throat of the thrust chamber and film cooled the hot gas side of the chamber wall without resulting in catastrophic failure. However, more testing is necessary to demonstrate that a catastropic failure would not occur if cracks developed further upstream between the injector and the throat, where the boundary layer has not been established. Since under normal conditions cracks are expected to form in the throat region of the thrust chamber, cracks must be initiated artificially in order to control their location. Several methods of crack initiation are discussed here.

  10. Activities to support the liquefied gaseous fuels spill test facility program. Final report

    SciTech Connect

    Sheesley, D.; King, S.B.; Routh, T.

    1997-03-01

    Approximately a hundred years ago the petrochemical industry was in its infancy, while the chemical industry was already well established. Today, both of these industries, which are almost indistinguishable, are a substantial part of the makeup of the U.S. economy and the lifestyle we enjoy. It is difficult to identify a single segment of our daily lives that isn`t affected by these industries and the products or services they make available for our use. Their survival and continued function in a competitive world market are necessary to maintain our current standard of living. The occurrence of accidents in these industries has two obvious effects: (1) the loss of product during the accident and future productivity because of loss of a portion of a facility or transport medium, and (2) the potential loss of life or injury to individuals, whether workers, emergency responders, or members of the general public. A great deal of work has been conducted at the Liquefied Gaseous Fuels Spill test Facility (LGFSTF) on hazardous spills. WRI has conducted accident investigations as well as provided information on the research results via the internet and bibliographies.

  11. Design and reliability optimization of a MEMS micro-hotplate for combustion of gaseous fuel

    SciTech Connect

    Manginell, R. P.

    2012-03-01

    This report will detail the process by which the silicon carbide (SiC) microhotplate devices, manufactured by GE, were imaged using IR microscopy equipment available at Sandia. The images taken were used as inputs to a finite element modeling (FEM) process using the ANSYS software package. The primary goal of this effort was to determine a method to measure the temperature of the microhotplate. Prior attempts to monitor the device's temperature by measuring its resistance had proven to be unreliable due to the nonlinearity of the doped SiC's resistance with temperature. As a result of this thermal modeling and IR imaging, a number of design recommendations were made to facilitate this temperature measurement. The lower heating value (LHV) of gaseous fuels can be measured with a catalyst-coated microhotplate calorimeter. GE created a silicon carbide (SiC) based microhotplate to address high-temperature survivability requirements for the application. The primary goal of this effort was to determine a method to measure the temperature of the microhotplate. Prior attempts to monitor the device's temperature by measuring its resistance had proven to be unreliable due to the non-linearity of the doped SiC's resistance with temperature. In this work, thermal modeling and IR imaging were utilized to determine the operation temperature as a function of parameters such as operation voltage and device sheet resistance. A number of design recommendations were made according to this work.

  12. Corrosion behavior of stainless steel in solid oxide fuel cell simulated gaseous environment

    SciTech Connect

    Ziomek-Moroz, M.; Covino, Bernard S., Jr.; Holcomb, Gordon R.; Cramer, Stephen D.; Matthes, Steven A.; Bullard, Sophie J.; Dunning, John S.; Alman, David E.; Wilson, Rick D.; Singh, P.

    2003-01-01

    Significant progress in reducing the operating temperature of solid oxide fuel cells (SOFC) from {approx}1000 C to {approx} 750 C may permit the replacement of currently used ceramic interconnects by metallic interconnects in planar SOFCs (PSOFC). The use of metallic interconnects will result in a substantial cost reduction of PSOFCs. The interconnects operate in severe gaseous environments, in which one side of the interconnect can be exposed to hydrogen and the other side to air or oxygen at temperatures up to 800 C. Similar environmental conditions can exist in devices used for separating hydrogen from CO after reforming methane and steam. Type 304 stainless steel was selected for this base line study aimed at understanding corrosion processes in dual gas environments. This paper discusses the oxidation resistance of 304 stainless steel exposed to a dual environment gas at 800 C. The dual environment consisted of air on one side of the specimen and 1% hydrogen in nitrogen on the other side. The surface characterization techniques used in this study were optical and scanning electron microscopy, as well as various x-ray techniques.

  13. Fuel decomposition and boundary-layer combustion processes of hybrid rocket motors

    NASA Technical Reports Server (NTRS)

    Chiaverini, Martin J.; Harting, George C.; Lu, Yeu-Cherng; Kuo, Kenneth K.; Serin, Nadir; Johnson, David K.

    1995-01-01

    Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated Polybutadiene) fuel cross-linked with diisocyanate was burned with GOX under various operating conditions. Large-amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed-line system and combustion chamber, the pressure oscillations were drastically reduced from +/-20% of the localized mean pressure to an acceptable range of +/-1.5% Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations are thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real-time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thickness burned and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented.

  14. Theoretical Rocket Performance of JP-4 Fuel with Several Fluorine-Oxygen Mixtures Assuming Equilibrium Composition

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford

    1958-01-01

    Theoretical rocket performance for equilibrium composition during expansion was calculated for JP-4 fuel with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, and equilibrium gas compositions. A correlation is given for the effect of chamber pressure on several of the parameters. The maximum value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827 atm) and an exit pressure of 1 atmosphere is 325.7 for 70.37 percent fluorine in the oxidant as compared with 284.9 and 305.1 for 100 percent oxygen and 100 percent fluorine, respectively.

  15. Fuel/oxidizer-rich high-pressure preburners. [staged-combustion rocket engine

    NASA Technical Reports Server (NTRS)

    Schoenman, L.

    1981-01-01

    The analyses, designs, fabrication, and cold-flow acceptance testing of LOX/RP-1 preburner components required for a high-pressure staged-combustion rocket engine are discussed. Separate designs of injectors, combustion chambers, turbine simulators, and hot-gas mixing devices are provided for fuel-rich and oxidizer-rich operation. The fuel-rich design addresses the problem of non-equilibrium LOX/RP-1 combustion. The development and use of a pseudo-kinetic combustion model for predicting operating efficiency, physical properties of the combustion products, and the potential for generating solid carbon is presented. The oxygen-rich design addresses the design criteria for the prevention of metal ignition. This is accomplished by the selection of materials and the generation of well-mixed gases. The combining of unique propellant injector element designs with secondary mixing devices is predicted to be the best approach.

  16. Hybrid rocket performance

    NASA Technical Reports Server (NTRS)

    Frederick, Robert A., Jr.

    1992-01-01

    A hybrid rocket is a system consisting of a solid fuel grain and a gaseous or liquid oxidizer. Figure 1 shows three popular hybrid propulsion cycles that are under current consideration. NASA MSFC has teamed with industry to test two hybrid propulsion systems that will allow scaling to motors of potential interest for Titan and Atlas systems, as well as encompassing the range of interest for SEI lunar ascent stages and National Launch System Cargo Transfer Vehicle (NLS CTV) and NLS deorbit systems. Hybrid systems also offer advantages as moderate-cost, environmentally acceptable propulsion system. The objective of this work was to recommend a performance prediction methodology for hybrid rocket motors. The scope included completion of: a literature review, a general methodology, and a simplified performance model.

  17. Modifying woody plants for efficient conversion to liquid and gaseous fuels

    SciTech Connect

    Dinus, R.J.; Dimmel, D.R.; Feirer, R.P.; Johnson, M.A.; Malcolm, E.W. )

    1990-07-01

    The Short Rotation Woody Crop Program (SRWCP), Department of Energy, is developing woody plant species as sources of renewable energy. Much progress has been made in identifying useful species, and testing site adaptability, stand densities, coppicing abilities, rotation lengths, and harvesting systems. Conventional plant breeding and intensive cultural practices have been used to increase above-ground biomass yields. Given these and foreseeable accomplishments, program leaders are now shifting attention to prospects for altering biomass physical and chemical characteristics, and to ways for improving the efficiency with which biomass can be converted to gaseous and liquid fuels. This report provides a review and synthesis of literature concerning the quantity and quality of such characteristics and constituents, and opportunities for manipulating them via conventional selection and breeding and/or molecular biology. Species now used by SRWCP are emphasized, with supporting information drawn from others as needed. Little information was found on silver maple (Acer saccharinum), but general comparisons (Isenberg 1981) suggest composition and behavior similar to those of the other species. Where possible, conclusions concerning means for and feasibility of manipulation are given, along with expected impacts on conversion efficiency. Information is also provided on relationships to other traits, genotype X environment interactions, and potential trade-offs or limitations. Biomass productivity per se is not addressed, except in terms of effects that may by caused by changes in constituent quality and/or quantity. Such effects are noted to the extent they are known or can be estimated. Likely impacts of changes, however effected, on suitability or other uses, e.g., pulp and paper manufacture, are notes. 311 refs., 4 figs., 9 tabs.

  18. Turbulent Mixing and Combustion of Multi-Phase Reacting Flows in Ramjet and Ducted Rocket Environment.

    DTIC Science & Technology

    1982-10-01

    both the fuel characteristics (gaseous fuels, boron-laden gaseous fuels, liquid fuels, and slurry fuels) and the flow field (axisymmetric-coaxial...D-A133 802 TURBULENT M IXING AND COMBUSTION OF MULTI-PHASE REACTING i/1 FLOWS I N RAMJET*RA U) NAVAL WEAPONS CENTER CHINA LAKE CA M J LEE ET AL. OCT...AFOSR MIPR 82-00010 .w~ TURBULENT MIXING AND COMBUSTION OF MULTI-PHASE REACTING FLOWS IN RAMJET AND DUCTED ROCKET ENVIRONMENT M. J. Lee K.C. Schadow

  19. Adsorption and chemical reaction of gaseous mixtures of hydrogen chloride and water on aluminum oxide and application to solid-propellant rocket exhaust clouds

    NASA Technical Reports Server (NTRS)

    Cofer, W. R., III; Pellett, G. L.

    1978-01-01

    Hydrogen chloride (HCl) and aluminum oxide (Al2O3) are major exhaust products of solid rocket motors (SRM). Samples of calcination-produced alumina were exposed to continuously flowing mixtures of gaseous HCl/H2O in nitrogen. Transient sorption rates, as well as maximum sorptive capacities, were found to be largely controlled by specific surface area for samples of alpha, theta, and gamma alumina. Sorption rates for small samples were characterized linearly with an empirical relationship that accounted for specific area and logarithmic time. Chemisorption occurred on all aluminas studied and appeared to form from the sorption of about a 2/5 HCl-to-H2O mole ratio. The chemisorbed phase was predominantly water soluble, yielding chloride/aluminum III ion mole ratios of about 3.3/1 suggestive of dissolved surface chlorides and/or oxychlorides. Isopiestic experiments in hydrochloric acid indicated that dissolution of alumina led to an increase in water-vapor pressure. Dissolution in aqueous SRM acid aerosol droplets, therefore, might be expected to promote evaporation.

  20. A feasibility study on using inkjet technology, micropumps, and MEMs as fuel injectors for bipropellant rocket engines

    NASA Astrophysics Data System (ADS)

    Glynne-Jones, Peter; Coletti, M.; White, N. M.; Gabriel, S. B.; Bramanti, C.

    2010-07-01

    Control over drop size distributions, injection rates, and geometrical distribution of fuel and oxidizer sprays in bi-propellant rocket engines has the potential to produce more efficient, more stable, less polluting rocket engines. This control also offers the potential of an engine that can be throttled, working efficiently over a wide range of output thrusts. Inkjet printing technologies, MEMS fuel atomizers, and piezoelectric injectors similar in concept to those used in diesel engines are considered for their potential to yield a new, more active injection scheme for a rocket engine. Inkjets are found to be unable to pump at sufficient pressures, and have possibly dangerous failure modes. Active injection is found to be feasible if high pressure drop along the injector plate is used. A conceptual design is presented and its basic behavior assessed.

  1. High temperature reformation of aluminum and chlorine compounds behind the Mach disk of a solid-fuel rocket exhaust

    NASA Technical Reports Server (NTRS)

    Park, C.

    1976-01-01

    Chemical reactions expected to occur among the constituents of solid-fuel rocket engine effluents in the hot region behind a Mach disk are analyzed theoretically. With the use of a rocket plume model that assumes the flow to be separated in the base region, and a chemical reaction scheme that includes evaporation of alumina and the associated reactions of 17 gas species, the reformation of the effluent is calculated. It is shown that AlClO and AlOH are produced in exchange for a corresponding reduction in the amounts of HCl and Al2O3. For the case of the space shuttle booster engines, up to 2% of the original mass of the rocket fuel can possibly be converted to these two new species and deposited in the atmosphere between the altitudes of 10 and 40 km. No adverse effects on the atmospheric environment are anticipated with the addition of these two new species.

  2. Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

    NASA Technical Reports Server (NTRS)

    Fortini, Anthony; Hendrix, Charles D.; Huff, Vearl N.

    1959-01-01

    The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

  3. Nuclear Cryogenic Propulsion Stage (NCPS) Fuel Element Testing in the Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    NASA Technical Reports Server (NTRS)

    Emrich, William J., Jr.

    2017-01-01

    To satisfy the Nuclear Cryogenic Propulsion Stage (NCPS) testing milestone, a graphite composite fuel element using a uranium simulant was received from the Oakridge National Lab and tested in the Nuclear Thermal Rocket Element Environmental Simulator (NTREES) at various operating conditions. The nominal operating conditions required to satisfy the milestone consisted of running the fuel element for a few minutes at a temperature of at least 2000 K with flowing hydrogen. This milestone test was successfully accomplished without incident.

  4. Investigation on the gaseous and particulate emissions of a compression ignition engine fueled with diesel-dimethyl carbonate blends.

    PubMed

    Cheung, C S; Zhu, Ruijun; Huang, Zuohua

    2011-01-01

    The effect of dimethyl carbonate (DMC) on the gaseous and particulate emissions of a diesel engine was investigated using Euro V diesel fuel blended with different proportions of DMC. Combustion analysis shows that, with the blended fuel, the ignition delay and the heat release rate in the premixed combustion phase increase, while the total combustion duration and the fuel consumed in the diffusion combustion phase decrease. Compared with diesel fuel, with an increase of DMC in the blended fuel, the brake thermal efficiency is slightly improved but the brake specific fuel consumption increases. On the emission side, CO increases significantly at low engine load but decreases at high engine load while HC decreases slightly. NO(x) reduces slightly but the reduction is not statistically significant, while NO(2) increases slightly. Particulate mass and number concentrations decrease upon using the blended fuel while the geometric mean diameter of the particles shifts towards smaller size. Overall speaking, diesel-DMC blends lead to significant improvement in particulate emissions while the impact on CO, HC and NO(x) emissions is small.

  5. High-Energy Propellant Rocket Firing at the Rocket Lab

    NASA Image and Video Library

    1955-01-21

    A rocket using high-energy propellant is fired from the Rocket Laboratory at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Rocket Lab was a collection of ten one-story cinderblock test cells located behind earthen barriers at the western edge of the campus. The rocket engines tested there were comparatively small, but the Lewis researchers were able to study different configurations, combustion performance, and injectors and nozzle design. The rockets were generally mounted horizontally and fired, as seen in this photograph of Test Cell No. 22. A group of fuels researchers at Lewis refocused their efforts after World War II in order to explore high energy propellants, combustion, and cooling. Research in these three areas began in 1945 and continued through the 1960s. The group of rocket researches was not elevated to a division branch until 1952. The early NACA Lewis work led to the development of liquid hydrogen as a viable propellant in the late 1950s. Following the 1949 reorganization of the research divisions, the rocket group began working with high-energy propellants such as diborane, pentaborane, and hydrogen. The lightweight fuels offered high levels of energy but were difficult to handle and required large tanks. In late 1954, Lewis researchers studied the combustion characteristics of gaseous hydrogen in a turbojet combustor. Despite poor mixing of the fuel and air, it was found that the hydrogen yielded more than a 90-percent efficiency. Liquid hydrogen became the focus of Lewis researchers for the next 15 years.

  6. Improving of Hybrid Rocket Engine on the Basis of Optimizing Design Fuel Grain

    NASA Astrophysics Data System (ADS)

    Oriekov, K. M.; Ushkin, M. P.

    2015-09-01

    This article examines the processes intrachamber in hybrid rocket engine (HRE) and the comparative assessment of the use of solid rocket motors (SRM) and HRE for meteorological rockets with a mass of payload of the 364 kg. Results of the research showed the possibility of a significant increase in the ballistic effectiveness of meteorological rocket.

  7. A solution to the problem of optimizing the fuel bias for a liquid propellant rocket by an application of the central limit theorem

    NASA Technical Reports Server (NTRS)

    Viera, W. J.

    1974-01-01

    A method of determining the fuel bias for a bipropellant liquid rocket that minimizes outage associated penalties on payload potential is presented. A fuel bias so derived is normally called the optimum fuel bias. The subjects discussed are: (1) probability density function of outage, (2) computer program listing, and (3) choosing the optimum fuel bias.

  8. Engineers demonstrate the pocket rocket

    NASA Technical Reports Server (NTRS)

    1995-01-01

    Part of Stennis Space Center's mission with its traveling exhibits is to educate the younger generation on how propulsion systems work. A popular tool is the 'pocket rocket,' which demonstrates how a hybrid rocket works. A hybrid rocket is a cross breed between a solid fuel rocket and a liquid fuel rocket.

  9. Engineers demonstrate the pocket rocket

    NASA Technical Reports Server (NTRS)

    1996-01-01

    Part of Stennis Space Center's mission with its traveling exhibits is to educate the younger generation on how propulsion systems work. A popular tool is the 'pocket rocket,' which demonstrates how a hybrid rocket works. A hybrid rocket is a cross breed between a solid fuel rocket and a liquid fuel rocket.

  10. Particulate and gaseous emissions from different wood fuels during combustion in a small-scale biomass heating system

    NASA Astrophysics Data System (ADS)

    Olave, R. J.; Forbes, E. G. A.; Johnston, C. R.; Relf, J.

    2017-05-01

    Woodchip is widely used as fuel in dedicated biomass and, even in some conventional energy generation plants. However, there are concerns about atmospheric air pollution from flue gases emitted during wood biomass combustion, particularly oxides of nitrogen (NOx) and particulates <10 μm diameter (PM10). In the United Kingdom (UK) a small scale biomass heat generation support scheme, the Renewable Heat Incentive (RHI), has been introduced. Qualifying criteria for this scheme have included limits for flue gas emissions of NOX and PM10 of 150 and 30 g per gigajoule (g/GJ) of energy input, respectively. In an experiment, three locally available types of Willow (Salix spp) and one of Sitka spruce (Picea sitchensis) woodchips, showed significant differences in physical and chemical constituents, gaseous and particulate emissions. During combustion in a 120 kW biomass system, air flows, flue gas temperatures and energy output correlated with gaseous emissions, NOx with raw fuel ash, nitrogen, phosphorus and potassium content, as did all flue gas particulate fractions. PM10 ranged from 30.3 to 105.7 g/GJ and NOx from 91.2 to 174.3 g/GJ. Sitka spruce produced significantly lower emissions of PM10 and NOx (27.5 and 52.6% less, respectively) than the three willow fuels, from which emissions were above the RHI emissions limits.

  11. Gaseous Surrogate Hydrocarbons for a Hifire Scramjet that Mimic Opposed Jet Extinction Limits for Cracked JP Fuels

    NASA Technical Reports Server (NTRS)

    Pellett, Gerald L.; Vaden, Sarah N.; Wilson, Lloyd G.

    2008-01-01

    This paper describes, first, the top-down methodology used to define simple gaseous surrogate hydrocarbon (HC) fuel mixtures for a hypersonic scramjet combustion subtask of the HiFIRE program. It then presents new and updated Opposed Jet Burner (OJB) extinction-limit Flame Strength (FS) data obtained from laminar non-premixed HC vs. air counterflow diffusion flames at 1-atm, which follow from earlier investigations. FS represents a strain-induced extinction limit based on cross-section-average air jet velocity, U(sub air), that sustains combustion of a counter jet of gaseous fuel just before extinction. FS uniquely characterizes a kinetically limited fuel combustion rate. More generally, Applied Stress Rates (ASRs) at extinction (U(sub air) normalized by nozzle or tube diameter, D(sub n or t) can directly be compared with extinction limits determined numerically using either a 1-D or (preferably) a 2-D Navier Stokes simulation with detailed transport and finite rate chemistry. The FS results help to characterize and define three candidate surrogate HC fuel mixtures that exhibit a common FS 70% greater than for vaporized JP-7 fuel. These include a binary fuel mixture of 64% ethylene + 36% methane, which is our primary recommendation. It is intended to mimic the critical flameholding limit of a thermally- or catalytically-cracked JP-7 like fuel in HiFIRE scramjet combustion tests. Our supporting experimental results include: (1) An idealized kinetically-limited ASR reactivity scale, which represents maximum strength non-premixed flames for several gaseous and vaporized liquid HCs; (2) FS characterizations of Colket and Spadaccini s suggested ternary surrogate, of 60% ethylene + 30% methane + 10% n-heptane, which matches the ignition delay of a typical cracked JP fuel; (3) Data showing how our recommended binary surrogate, of 64% ethylene + 36% methane, has an identical FS; (4) Data that characterize an alternate surrogate of 44% ethylene + 56% ethane with identical

  12. DNS of moderate-temperature gaseous mixing layers laden with multicomponent-fuel drops

    NASA Technical Reports Server (NTRS)

    Clercq, P. C. Le; Bellan, J.

    2004-01-01

    A formulation representing multicomponent-fuel (MC-fuel) composition as a Probability Distribution Function (PDF) depending on the molar weight is used to construct a model of a large number of MC-fuel drops evaporating in a gas flow, so as to assess the extent of fuel specificity on the vapor composition.

  13. DNS of moderate-temperature gaseous mixing layers laden with multicomponent-fuel drops

    NASA Technical Reports Server (NTRS)

    Clercq, P. C. Le; Bellan, J.

    2004-01-01

    A formulation representing multicomponent-fuel (MC-fuel) composition as a Probability Distribution Function (PDF) depending on the molar weight is used to construct a model of a large number of MC-fuel drops evaporating in a gas flow, so as to assess the extent of fuel specificity on the vapor composition.

  14. Critical Design Parameters for Pylon-Aided Gaseous Fuel Injection Upstream of a Flameholding Cavity

    DTIC Science & Technology

    2009-03-01

    cavity depth (in) FPP flammable plume percentage h pylon height (in) penetration of plume core (in) total fuel plume penetration (in) l pylon...metrics include: mixing efficiency (η), fuel plume area (Ap), flammable fuel plume area (Af) , flammable fuel plume percentage ( FPP ), species...combustion. Another key metric is the flammable plume percentage ( FPP ), as seen in Equation 10. This metric represents the percentage of the fuel

  15. Effect of buoyancy on fuel containment in an open-cycle gas-core nuclear rocket engine.

    NASA Technical Reports Server (NTRS)

    Putre, H. A.

    1971-01-01

    Analysis aimed at determining the scaling laws for the buoyancy effect on fuel containment in an open-cycle gas-core nuclear rocket engine, so conducted that experimental conditions can be related to engine conditions. The fuel volume fraction in a short coaxial flow cavity is calculated with a programmed numerical solution of the steady Navier-Stokes equations for isothermal, variable density fluid mixing. A dimensionless parameter B, called the Buoyancy number, was found to correlate the fuel volume fraction for large accelerations and various density ratios. This parameter has the value B = 0 for zero acceleration, and B = 350 for typical engine conditions.

  16. Raman Gas Species Measurements in Hydrocarbon-Fueled Rocket Engine Injector Flows

    NASA Technical Reports Server (NTRS)

    Wehrmeyer, Joseph; Hartfield, Roy J., Jr.; Trinh, Huu P.; Dobson, Chris C.; Eskridge, Richard H.

    2000-01-01

    Rocket engine propellent injector development at NASA-Marshall includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellent mass transported to Mars for future manned Mars missions. The Raman technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented, as well as a high pressure demonstration in the NASA-Marshall Modular Combustion Test Artice, using the liquid methane-liquid oxygen propellant system

  17. Fundamental phenomena on fuel decomposition and boundary-layer combustion processes with applications to hybrid rocket motors

    NASA Technical Reports Server (NTRS)

    Kuo, Kenneth K.; Lu, Yeu-Cherng; Chiaverini, Martin J.; Harting, George C.; Johnson, David K.; Serin, Nadir

    1995-01-01

    The experimental study on the fundamental processes involved in fuel decomposition and boundary-layer combustion in hybrid rocket motors is continuously being conducted at the High Pressure Combustion Laboratory of The Pennsylvania State University. This research will provide a useful engineering technology base in the development of hybrid rocket motors as well as a fundamental understanding of the complex processes involved in hybrid propulsion. A high-pressure, 2-D slab motor has been designed, manufactured, and utilized for conducting seven test firings using HTPB fuel processed at PSU. A total of 20 fuel slabs have been received from the Mcdonnell Douglas Aerospace Corporation. Ten of these fuel slabs contain an array of fine-wire thermocouples for measuring solid fuel surface and subsurface temperatures. Diagnostic instrumentation used in the test include high-frequency pressure transducers for measuring static and dynamic motor pressures and fine-wire thermocouples for measuring solid fuel surface and subsurface temperatures. The ultrasonic pulse-echo technique as well as a real-time x-ray radiography system have been used to obtain independent measurements of instantaneous solid fuel regression rates.

  18. Response of selected plant and insect species to simulated solid rocket exhaust mixtures and to exhaust components from solid rocket fuels

    NASA Technical Reports Server (NTRS)

    Heck, W. W.; Knott, W. M.; Stahel, E. P.; Ambrose, J. T.; Mccrimmon, J. N.; Engle, M.; Romanow, L. A.; Sawyer, A. G.; Tyson, J. D.

    1980-01-01

    The effects of solid rocket fuel (SRF) exhaust on selected plant and and insect species in the Merritt Island, Florida area was investigated in order to determine if the exhaust clouds generated by shuttle launches would adversely affect the native, plants of the Merritt Island Wildlife Refuge, the citrus production, or the beekeeping industry of the island. Conditions were simulated in greenhouse exposure chambers and field chambers constructed to model the ideal continuous stirred tank reactor. A plant exposure system was developed for dispensing and monitoring the two major chemicals in SRF exhaust, HCl and Al203, and for dispensing and monitoring SRF exhaust (controlled fuel burns). Plants native to Merritt Island, Florida were grown and used as test species. Dose-response relationships were determined for short term exposure of selected plant species to HCl, Al203, and mixtures of the two to SRF exhaust.

  19. Results of Labscale Hybrid Rocket Motor investigation

    NASA Technical Reports Server (NTRS)

    Greiner, B.; Frederick, R. A., Jr.

    1992-01-01

    This work was performed to establish a labscale hybrid rocket motor test and evaluation capability at NASA Marshall Space Flight Center. The scope included activation of a Labscale Hybrid Motor, determination of baseline burning rates for PMMA fuel, and replication of pressure oscillations for HTPB fuel. The 0.820-in.-diam port, 10-in.-long fuel grains were burned for two seconds with gaseous oxygen. PMMA fuels were tested at oxygen fluxes from 0.047 lbm/sec sq in. to 0.378 lbm/sec sq in., and the HTPB fuel was evaluated at 0.378 lbm/sec sq in. The results showed that the labscale hybrid motor replicated previously reported PMMA fuel regression rates. The results also replicated low-frequency (less than 100 Hz) pressure oscillations that have been observed for HTPB fuels. These results establish the Labscale Hybrid Motor facility at MSFC.

  20. Results of Labscale Hybrid Rocket Motor investigation

    NASA Technical Reports Server (NTRS)

    Greiner, B.; Frederick, R. A., Jr.

    1992-01-01

    This work was performed to establish a labscale hybrid rocket motor test and evaluation capability at NASA Marshall Space Flight Center. The scope included activation of a Labscale Hybrid Motor, determination of baseline burning rates for PMMA fuel, and replication of pressure oscillations for HTPB fuel. The 0.820-in.-diam port, 10-in.-long fuel grains were burned for two seconds with gaseous oxygen. PMMA fuels were tested at oxygen fluxes from 0.047 lbm/sec sq in. to 0.378 lbm/sec sq in., and the HTPB fuel was evaluated at 0.378 lbm/sec sq in. The results showed that the labscale hybrid motor replicated previously reported PMMA fuel regression rates. The results also replicated low-frequency (less than 100 Hz) pressure oscillations that have been observed for HTPB fuels. These results establish the Labscale Hybrid Motor facility at MSFC.

  1. Plasma core nuclear rocket technology

    NASA Astrophysics Data System (ADS)

    Latham, Thomas S.; Roman, Ward C.; Johnson, Bruce V.

    1993-06-01

    The nuclear lightbulb (NLB) rocket propulsion concept furnishes specific impulse above 2000 sec in conjunction with the greater-than-50,000 lb thrust levels required for rapid transit-time round-trip Mars missions requiring low initial mass in earth orbit. The NLB transfers energy from the gaseous nuclear fuel region to a hydrogen propellant via thermal radiation, thereby precluding material temperature constraints. An evaluation is presently made of technology and test method readiness for implementation and validation of this propulsion system concept.

  2. Plasma core nuclear rocket technology

    SciTech Connect

    Latham, T.S.; Roman, W.C.; Johnson, B.V.

    1993-06-01

    The nuclear lightbulb (NLB) rocket propulsion concept furnishes specific impulse above 2000 sec in conjunction with the greater-than-50,000 lb thrust levels required for rapid transit-time round-trip Mars missions requiring low initial mass in earth orbit. The NLB transfers energy from the gaseous nuclear fuel region to a hydrogen propellant via thermal radiation, thereby precluding material temperature constraints. An evaluation is presently made of technology and test method readiness for implementation and validation of this propulsion system concept. 13 refs.

  3. Impact of the 0.1% fuel sulfur content limit in SECA on particle and gaseous emissions from marine vessels

    NASA Astrophysics Data System (ADS)

    Zetterdahl, Maria; Moldanová, Jana; Pei, Xiangyu; Pathak, Ravi Kant; Demirdjian, Benjamin

    2016-11-01

    Emissions were measured on-board a ship in the Baltic Sea, which is a sulfur emission control area (SECA), before and after the implementation of the strict fuel sulfur content (FSC) limit of 0.1 m/m% S on the 1st of January 2015. Prior to January 2015, the ship used a heavy fuel oil (HFO) but switched to a low-sulfur residual marine fuel oil (RMB30) after the implementation of the new FSC limit. The emitted particulate matter (PM) was measured in terms of mass, number, size distribution, volatility, elemental composition, content of organics, black and elemental carbon, polycyclic aromatic hydrocarbons (PAHs), microstructure and micro-composition, along with the gaseous emissions at different operating conditions. The fuel change reduced emissions of PM mass up to 67%. The number of particles emitted remained unchanged and were dominated by nanoparticles. Furthermore, the fuel change resulted in an 80% reduction of SO2 emissions and decreased emissions of total volatile organic compounds (VOCs). The emissions of both monoaromatic and lighter polyaromatic hydrocarbon compounds increased with RMB30, while the heavy, PM-bound PAH species that belong to the carcinogenic PAH family were reduced. Emissions of BC remained similar between the two fuels. This study indicates that the use of low-sulfur residual marine fuel oil is a way to comply with the new FSC regulation and will reduce the anthropogenic load of SO2 emissions and secondary PM formed from SO2. Emissions of primary particles, however, remain unchanged and do not decrease as much as would be expected if distilled fuel was used. This applies both to the number of particles emitted and some toxic components, such as heavy metals, PAHs or elemental carbon (EC). The micro-composition analyses showed that the soot particles emitted from RMB30 combustion often do not have any trace of sulfur compared with particles from HFO combustion, which always have a sulfur content over 1%m/m. The soot sulfur content can

  4. Program ELM: A tool for rapid thermal-hydraulic analysis of solid-core nuclear rocket fuel elements

    NASA Technical Reports Server (NTRS)

    Walton, James T.

    1992-01-01

    This report reviews the state of the art of thermal-hydraulic analysis codes and presents a new code, Program ELM, for analysis of fuel elements. ELM is a concise computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in a nuclear thermal rocket reactor with axial coolant passages. The program was developed as a tool to swiftly evaluate various heat transfer coefficient and friction factor correlations generated for turbulent pipe flow with heat addition which have been used in previous programs. Thus, a consistent comparison of these correlations was performed, as well as a comparison with data from the NRX reactor experiments from the Nuclear Engine for Rocket Vehicle Applications (NERVA) project. This report describes the ELM Program algorithm, input/output, and validation efforts and provides a listing of the code.

  5. Nuclear Cryogenic Propulsion Stage (NCPS) Fuel Element Testing in the Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    NASA Technical Reports Server (NTRS)

    Emrich, William J., Jr.

    2017-01-01

    To support the on-going nuclear thermal propulsion effort, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The facility to perform this testing is referred to as the Nuclear Thermal Rocket Element Environment Simulator (NTREES). Last year NTREES was successfully used to satisfy a testing milestone for the Nuclear Cryogenic Propulsion Stage (NCPS) project and met or exceeded all required objectives.

  6. Development of a model for baffle energy dissipation in liquid fueled rocket engines

    NASA Astrophysics Data System (ADS)

    Miller, Nathan A.

    In this thesis the energy dissipation from a combined hub and blade baffle structure in a combustion chamber of a liquid-fueled rocket engine is modeled and computed. An analytical model of the flow stabilization due to the effect of combined radial and hub blades was developed. The rate of energy dissipation of the baffle blades was computed using a corner-flow model that included unsteady flow separation and turbulence effects. For the inviscid portion of the flow field, a solution methodology was formulated using an eigenfunction expansion and a velocity potential matching technique. Parameters such as local velocity, elemental path length, effective viscosity, and local energy dissipation rate were computed as a function of the local angle alpha for a representative baffle blade, and compared to results predicted by the Baer-Mitchell blade dissipation model. The sensitivity of the model to the overall engine acoustic oscillation mode, blade length, and thickness was also computed and compared to previous results. Additional studies were performed to determine the sensitivity to input parameters such as the dimensionless turbulence coefficient, the location of the potential difference in the generation of the dividing streamline, the number of baffle blades and the size of the central hub. Stability computations of a test engine indicated that when the baffle length is increased, the baffles provide increased stabilization effects. The model predicts greatest dissipation for radial modes with a hub radius at approximately half the chamber's radius.

  7. Test program to provide confidence in liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

    NASA Technical Reports Server (NTRS)

    Armstrong, E. S.

    1986-01-01

    An experimental program has been planned at the NASA Lewis Research Center to build confidence in the feasibility of liquid oxygen cooling for hydrocarbon fueled rocket engines. Although liquid oxygen cooling has previously been incorporated in test hardware, more runtime is necessary to gain confidence in this concept. In the previous tests, small oxygen leaks developed at the throat of the thrust chamber and film cooled the hot-gas side of the chamber wall without resulting in catastrophic failure. However, more testing is necessary to demonstrate that a catastrophic failure would not occur if cracks developed further upstream between the injector and the throat, where the boundary layer has not been established. Since under normal conditions cracks are expected to form in the throat region of the thrust chamber, cracks must be initiated artificially in order to control their location. Several methods of crack initiation are discussed in this report. Four thrust chambers, three with cracks and one without, should be tested. The axial location of the cracks should be varied parametrically. Each chamber should be instrumented to determine the effects of the cracks, as well as the overall performance and durability of the chambers.

  8. A PEMS study of the emissions of gaseous pollutants and ultrafine particles from gasoline- and diesel-fueled vehicles

    NASA Astrophysics Data System (ADS)

    Huang, Cheng; Lou, Diming; Hu, Zhiyuan; Feng, Qian; Chen, Yiran; Chen, Changhong; Tan, Piqiang; Yao, Di

    2013-10-01

    On-road emission measurements of gasoline- and diesel-fueled vehicles were conducted by a portable emission measurement system (PEMS) in Shanghai, China. Horiba OBS 2200 and TSI EEPS 3090 were employed to detect gaseous and ultrafine particle emissions during the tests. The driving-based emission factors of gaseous pollutants and particle mass and number were obtained on various road types. The average NOx emission factors of the diesel bus, diesel car, and gasoline car were 8.86, 0.68, and 0.17 g km-1, all of which were in excess of their emission limits. The particle number emission factors were 7.06 × 1014, 6.08 × 1014, and 1.57 × 1014 km-1, generally higher than the results for similar vehicle types reported in the previous studies. The size distributions of the particles emitted from the diesel vehicles were mainly concentrated in the accumulation mode, while those emitted from the gasoline car were mainly distributed in the nucleation mode. Both gaseous and particle emission rates exhibit significant correlations with the change in vehicle speed and power demand. The lowest emission rates for each vehicle type were produced during idling. The highest emission rates for each vehicle type were generally found in high-VSP bins. The particle number emission rates of the gasoline car show the strongest growth trend with increasing VSP and speed. The particle number emission for the gasoline car increased by 3 orders of magnitude from idling to the highest VSP and driving speed conditions. High engine power caused by aggressive driving or heavy loads is the main contributor to high emissions for these vehicles in real-world situations.

  9. Biological production of liquid and gaseous fuels from coal synthesis gas

    SciTech Connect

    Antorrena, G.M.; Vega, J.L.; Clausen, E.C.; Gaddy, J.L.

    1988-01-01

    Cultures of microorganisms have been isolated that convert CO, H/sub 2/ and CO/sub 2/ in coal synthesis gas into methane or ethanol. The reactions are severely mass transfer limited and bioreactor design will be a critical factor in the application of this technology. This paper presents results of culture isolation studies and development of continuous reactors for these cultures. The results of bubble columns and stirred tank reactors are presented and discussed. Methods for defining mass transfer coefficients and intrinsic kinetics are presented. Operation of these gaseous fermentations at high pressure has enabled complete conversion in reaction times of a few minutes.

  10. Biological studies in the impact zone of the Liquefied Gaseous Fuels Spill Test Facility in Frenchman Flat, Nevada

    SciTech Connect

    Hunter, R.B.; Saethre, M.B.; Medica, P.A.; Greger, P.D.; Romney, E.M.

    1991-01-01

    Desert shrubs and rodents were monitored downwind of the Department of Energy Liquefied Gaseous Fuels Spill Test Facility (LGF), which is situated on a dry lake bed (playa). Plants were censused in 1981 and 1986 through 1990; rodent survival was measured from 1986 through 1990. During that time there were no apparent effects of the spill tests on animals or plants off the edge of the playa, which extends more than 2.5 kilometers from the facility. Plant populations increased in volume from 1981 through 1986, then declined precipitously during drought in 1989 and 1990. Rodent populations also declined during the drought. Some effects of spilled hydrogen fluoride gas were seen on plants growing on manmade mounds on the playa surface. Animal and bird species seen in the vicinity of the LGF are also reported. 11 refs., 10 figs., 16 tabs.

  11. Integrated Advanced Reciprocating Internal Combustion Engine System for Increased Utilization of Gaseous Opportunity Fuels

    SciTech Connect

    Pratapas, John; Zelepouga, Serguei; Gnatenko, Vitaliy; Saveliev, Alexei; Jangale, Vilas; Li, Hailin; Getz, Timothy; Mather, Daniel

    2013-08-31

    The project is addressing barriers to or opportunities for increasing distributed generation (DG)/combined heat and power (CHP) use in industrial applications using renewable/opportunity fuels. This project brings together novel gas quality sensor (GQS) technology with engine management for opportunity fuels such as landfill gas, digester gas and coal bed methane. By providing the capability for near real-time monitoring of the composition of these opportunity fuels, the GQS output can be used to improve the performance, increase efficiency, raise system reliability, and provide improved project economics and reduced emissions for engines used in distributed generation and combined heat and power.

  12. 70. VIEW OF FUEL APRON FROM EAST SIDE OF LAUNCH ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    70. VIEW OF FUEL APRON FROM EAST SIDE OF LAUNCH PAD. ROCKET FUEL TANKS ON LEFT; GASEOUS NITROGEN AND HELIUM TANKS IN CENTER; AND A LARGE LIQUID NITROGEN TANK ON RIGHT. SKID 1 FOR GASEOUS NITROGEN TRANSFER AND SKID 5 FOR HELIUM TRANSFER IN THE CENTER RIGHT PORTION OF THE PHOTOGRAPH. - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 East, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

  13. Emission factors of gaseous pollutants from recent kerosene space heaters and fuels available in France in 2010.

    PubMed

    Carteret, M; Pauwels, J-F; Hanoune, B

    2012-08-01

    Laboratory measurements of the gaseous emission factors (EF) from two recent kerosene space heaters (wick and injector) with five different fuels have been conducted in an 8-m(3) environmental chamber. The two heaters tested were found to emit mainly CO(2), CO, NO, NO(2), and some volatile organic compounds (VOCs). NO(2) is continuously emitted during use, with an EF of 100-450 μg per g of consumed fuel. CO is normally emitted mainly during the first minutes of use (up to 3 mg/g). Formaldehyde and benzene EFs were quantified at 15 and 16 μg/g, respectively, for the wick heater. Some other VOCs, such as 1,3-butadiene, were detected with lower EFs. We demonstrated the unsuitability of a 'biofuel' containing fatty acid methyl esters for use with the wick heater, and that the accumulation of soot on the same heater, whatever the fuel, leads to a dramatic increase in the CO EF, up to 16 mg/g, which could be responsible for chronic and acute CO intoxications. Our results show that in spite of new technologies and emission standards for unvented kerosene space heaters, as well as for the fuels, the use of these heaters in indoor environments still leads to NO(x) levels in excess of current health recommendations. Whereas injection heaters generate more nitrogen oxides than wick heaters, prolonged use of the latter leads to a soot buildup, concomitant with high CO emissions, which could be responsible for acute and chronic intoxications. The use of a biofuel in a wick heater is also of concern. Maintenance of the heaters and adequate ventilation of the room during use of kerosene space heaters are therefore of prime importance to reduce personal exposure. © 2011 John Wiley & Sons A/S.

  14. Liquefied Gaseous Fuels Safety and Environmental Control Assessment Program: second status report

    SciTech Connect

    1980-10-01

    Volume 2 consists of 19 reports describing technical effort performed by Government Contractors in the area of LNG Safety and Environmental Control. Report topics are: simulation of LNG vapor spread and dispersion by finite element methods; modeling of negatively buoyant vapor cloud dispersion; effect of humidity on the energy budget of a liquefied natural gas (LNG) vapor cloud; LNG fire and explosion phenomena research evaluation; modeling of laminar flames in mixtures of vaporized liquefied natural gas (LNG) and air; chemical kinetics in LNG detonations; effects of cellular structure on the behavior of gaseous detonation waves under transient conditions; computer simulation of combustion and fluid dynamics in two and three dimensions; LNG release prevention and control; the feasibility of methods and systems for reducing LNG tanker fire hazards; safety assessment of gelled LNG; and a four band differential radiometer for monitoring LNG vapors.

  15. Air emission from the co-combustion of alternative derived fuels within cement plants: Gaseous pollutants.

    PubMed

    Richards, Glen; Agranovski, Igor E

    2015-02-01

    Cement manufacturing is a resource- and energy-intensive industry, utilizing 9% of global industrial energy use while releasing more than 5% of global carbon dioxide (CO₂) emissions. With an increasing demand of production set to double by 2050, so too will be its carbon footprint. However, Australian cement plants have great potential for energy savings and emission reductions through the substitution of combustion fuels with a proportion of alternative derived fuels (ADFs), namely, fuels derived from wastes. This paper presents the environmental emissions monitoring of 10 cement batching plants while under baseline and ADF operating conditions, and an assessment of parameters influencing combustion. The experiential runs included the varied substitution rates of seven waste streams and the monitoring of seven target pollutants. The co-combustion tests of waste oil, wood chips, wood chips and plastic, waste solvents, and shredded tires were shown to have the minimal influence when compared to baseline runs, or had significantly reduced the unit mass emission factor of pollutants. With an increasing ADF% substitution, monitoring identified there to be no subsequent emission effects and that key process parameters contributing to contaminant suppression include (1) precalciner and kiln fuel firing rate and residence time; (2) preheater and precalciner gas and material temperature; (3) rotary kiln flame temperature; (4) fuel-air ratio and percentage of excess oxygen; and (5) the rate of meal feed and rate of clinker produced.

  16. Experimental Performance of Area Ratio 200, 25 and 8 Nozzles on JP-4 Fuel and Liquid Oxygen Rocket Engine

    NASA Technical Reports Server (NTRS)

    Lovell, J. Calvin; Samanich, Nick E.; Barnett, Donald O.

    1960-01-01

    The performance of an area ratio 200 bell-shaped nozzle, an area ratio 25 bell-shaped nozzle, and an area ratio 8 conic nozzle on a JP-4 fuel and liquid-oxygen rocket engine has been determined. Tests were conducted using a nominal 4000-pound-thrust rocket in the Lewis 10- by 10-foot supersonic tunnel, which provided the altitude environment needed for fully expanded nozzle flow. The area ratio 200 nozzle had a vacuum thrust coefficient of 1.96, compared with 1.82 and 1.70 for the area ratio 25 and 8 nozzles, respectively. These values are approximately equal to those for theoretical frozen expansion. The measured value of vacuum specific impulse for the area ratio 200 nozzle was 317 seconds for a combustion-chamber characteristic velocity of 5200 feet per second. The vacuum-specific-impulse increase for the area-ratio increase from 8 to 200 was 46 seconds.

  17. Combustion oscillation study in a kerosene fueled rocket-based combined-cycle engine combustor

    NASA Astrophysics Data System (ADS)

    Huang, Zhi-Wei; He, Guo-Qiang; Qin, Fei; Xue, Rui; Wei, Xiang-Geng; Shi, Lei

    2016-12-01

    This study reports the combustion oscillation features in a three-dimensional (3D) rocket-based combined-cycle (RBCC) engine combustor under flight Mach number (Mflight) 3.0 conditions both experimentally and numerically. Experiment is performed on a direct-connect ground test facility, which measures the wall pressure along the flow-path. High-speed imaging of the flame luminosity and schlieren is carried out at exit of the primary rocket. Compressible reactive large eddy simulation (LES) with reduced chemical kinetics of a surrogate model for kerosene is performed to further understand the combustion oscillation mechanisms in the combustor. LES results are validated with experimental data by the time-averaged and root mean square (RMS) pressure values, and show acceptable agreement. Effects of the primary rocket jet on pressure oscillation in the combustor are analyzed. Relation of the high speed rocket jet oscillation, which is thought to among the most probable sources of combustion oscillation, with the RBCC combustor is recognized. Results reveal that the unsteady over-expanded rocket jet has significant impacts on the combustion oscillation feature of the RBCC combustor, which is different from a thermo-acoustics type oscillation. The rocket jet/air inflow physical interactions under different rocket jet expansion degrees are experimentally studied.

  18. Preliminary assessment of systems for deriving liquid and gaseous fuels from waste or grown organics

    NASA Technical Reports Server (NTRS)

    Graham, R. W.; Reynolds, T. W.; Hsu, Y. Y.

    1976-01-01

    The overall feasibility of the chemical conversion of waste or grown organic matter to fuel is examined from the technical, economic, and social viewpoints. The energy contribution from a system that uses waste and grown organic feedstocks is estimated as 4 to 12 percent of our current energy consumption. Estimates of today's market prices for these fuels are included. Economic and social issues are as important as technology in determining the feasibility of such a proposal. An orderly program of development and demonstration is recommended to provide reliable data for an assessment of the viability of the proposal.

  19. Catalytic Microtube Rocket Igniter

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Deans, Matthew C.

    2011-01-01

    Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each

  20. Hazards Induced by Breach of Liquid Rocket Fuel Tanks: Conditions and Risks of Cryogenic Liquid Hydrogen-Oxygen Mixture Explosions

    NASA Technical Reports Server (NTRS)

    Osipov, Viatcheslav; Muratov, Cyrill; Hafiychuk, Halyna; Ponizovskya-Devine, Ekaterina; Smelyanskiy, Vadim; Mathias, Donovan; Lawrence, Scott; Werkheiser, Mary

    2011-01-01

    We analyze the data of purposeful rupture experiments with LOx and LH2 tanks, the Hydrogen-Oxygen Vertical Impact (HOVI) tests that were performed to clarify the ignition mechanisms, the explosive power of cryogenic H2/Ox mixtures under different conditions, and to elucidate the puzzling source of the initial formation of flames near the intertank section during the Challenger disaster. We carry out a physics-based analysis of general explosions scenarios for cryogenic gaseous H2/Ox mixtures and determine their realizability conditions, using the well-established simplified models from the detonation and deflagration theory. We study the features of aerosol H2/Ox mixture combustion and show, in particular, that aerosols intensify the deflagration flames and can induce detonation for any ignition mechanism. We propose a cavitation-induced mechanism of self-ignition of cryogenic H2/Ox mixtures that may be realized when gaseous H2 and Ox flows are mixed with a liquid Ox turbulent stream, as occurred in all HOVI tests. We present an overview of the HOVI tests to make conclusion on the risk of strong explosions in possible liquid rocket incidents and provide a semi-quantitative interpretation of the HOVI data based on aerosol combustion. We uncover the most dangerous situations and discuss the foreseeable risks which can arise in space missions and lead to tragic outcomes. Our analysis relates to only unconfined mixtures that are likely to arise as a result of liquid propellant space vehicle incidents.

  1. Microfabricated Liquid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Epstein, Alan H.; Joppin, C.; Kerrebrock, J. L.; Schneider, Steven J. (Technical Monitor)

    2003-01-01

    Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.

  2. Study on Gaseous Effluent Treatment for Dissolution Step of Spent Nuclear Fuel Reprocessing

    SciTech Connect

    Mineo, H.; Iizuka, M.; Fujisaki, S.; Hotoku, S.; Asakura, T.; Uchiyama, G.

    2002-02-27

    Behavior of radioiodine and carbon-14 during spent fuel dissolution was studied in a bench-scale reprocessing test rig where 29 and 44 GWdt-1 spent fuels were respectively dissolved. Decontamination factor of AGS (silica-gel impregnated with silver nitrate) column for iodine-129 removal was measured to be more than 36,000. The measurement of iodine-129 profile in the adsorption column showed that the nuclide was effectively trapped by the adsorbent. Measurement of iodine-129 in the dissolver solution after the iodine-stripping operation using NO2 gas at 363 K, revealed that less than 0.57% of total iodine-129 generated, which was estimated by ORIGEN II calculation, was remained in the dissolver solution. Also, measurement of iodine-129 by an iodine-stripping operation from the dissolver solution using potassium iodate showed that another 2.72% of total iodine-129 precipitated as iodide. In addition, about 70 % of total iodine generated was measured in the AGS columns. Rest of iodine-129 was supposed to adsorb to a HEPA filter and the inner surface of dissolver off-gas lines. Those results on iodine-129 distribution were found to be almost identical to the results obtained in the study using iodine-131 as tracer and the results reported by other works. It was demonstrated that the two-steps iodine-stripping method using potassium iodate could expel additional iodine from the solution, more effectively than iodine-stripping operation using NO2 gas. Iodine-131 was also detected on the AGS columns at the spent fuel dissolution. Increasing burnup showed larger amount of iodine-131 since amount of curium-244 contained in the spent fuel increased with the burnup. Release of carbon-14 as carbon dioxide during dissolution was found to occur when the release of krypton-85. From the 14CO2 measurement, initial nitrogen-14 concentration in the fuel was estimated to be about several ppm, which was within the range reported.

  3. Identification of vapor-phase chemical warfare agent simulants and rocket fuels using laser-induced breakdown spectroscopy

    SciTech Connect

    Stearns, Jaime A.; McElman, Sarah E.; Dodd, James A.

    2010-05-01

    Application of laser-induced breakdown spectroscopy (LIBS) to the identification of security threats is a growing area of research. This work presents LIBS spectra of vapor-phase chemical warfare agent simulants and typical rocket fuels. A large dataset of spectra was acquired using a variety of gas mixtures and background pressures and processed using partial least squares analysis. The five compounds studied were identified with a 99% success rate by the best method. The temporal behavior of the emission lines as a function of chamber pressure and gas mixture was also investigated, revealing some interesting trends that merit further study.

  4. The application of near-infrared spectroscopy for the quality control analysis of rocket propellant fuel pre-mixes.

    PubMed

    Judge, Michael D

    2004-03-10

    The viability of near-infrared (NIR) spectroscopy as a technique for the quality control analysis of ingredient concentrations in a rocket propellant fuel liquid pre-mix was investigated. The pre-mix analyzed consisted of a polybutadiene pre-polymer, a plasticizer and two antioxidants. It was determined that NIR spectroscopy offered a fast and convenient method of verifying the percentage level of all four ingredients while requiring no sample preparation. The NIR methodology exhibited a high level of accuracy and precision. There was also a clear indication that the technique allowed monitoring of antioxidant depletion in the pre-mix on ageing.

  5. Starting of rocket engine at conditions of simulated altitude using crude monoethylaniline and other fuels with mixed acid

    NASA Technical Reports Server (NTRS)

    Ladanyi, Dezso J; Sloop, John L; Humphrey, Jack C; Morrell, Gerald

    1950-01-01

    Experiments were conducted at sea level and pressure altitude of about 55,000 feet at various temperatures to determine starting characteristics of a commercial rocket engine using crude monoethylaniline and other fuels with mixed acid. With crude monoethylaniline, ignition difficulties were encountered at temperatures below about 20 degrees F. With mixed butyl mercaptans, water-white turpentine, and x-pinene, no starting difficulties were experienced at temperatures as low as minus 74 degrees F. Turpentine and x-pinene, however, sometimes left deposits on the injector face. With blends containing furfuryl alcohol and with other blends, difficulties were experienced either from appreciable deposits or from starting.

  6. Combustion Processes in Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Venkateswaran,S.; Merkle, C. L.

    1996-01-01

    In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. Hybrid propulsion systems use a solid fuel such as hydroxyl-terminated polybutadiene (HTPB) along with a gaseous/liquid oxidizer. The performance of hybrid combustors depends on the convective and radiative heat fluxes to the fuel surface, the rate of pyrolysis in the solid phase, and the turbulent combustion processes in the gaseous phases. These processes in combination specify the regression rates of the fuel surface and thereby the utilization efficiency of the fuel. In this paper, we employ computational fluid dynamics (CFD) techniques in order to gain a quantitative understanding of the physical trends in hybrid rocket combustors. The computational modeling is tailored to ongoing experiments at Penn State that employ a two dimensional slab burner configuration. The coordinated computational/experimental effort enables model validation while providing an understanding of the experimental observations. Computations to date have included the full length geometry with and with the aft nozzle section as well as shorter length domains for extensive parametric characterization. HTPB is sed as the fuel with 1,3 butadiene being taken as the gaseous product of the pyrolysis. Pure gaseous oxygen is taken as the oxidizer. The fuel regression rate is specified using an Arrhenius rate reaction, which the fuel surface temperature is given by an energy balance involving gas-phase convection and radiation as well as thermal conduction in the solid-phase. For the gas-phase combustion, a two step global reaction is used. The standard kappa - epsilon model is used for turbulence closure. Radiation is presently treated using a simple diffusion approximation which is valid for large optical path lengths, representative of radiation from soot particles. Computational results are obtained to determine the trends in the fuel burning or

  7. Combustion of Gaseous Fuels with High Temperature Air in Normal- and Micro-gravity Conditions

    NASA Technical Reports Server (NTRS)

    Wang, Y.; Gupta, A. K.

    2001-01-01

    The objective of this study is determine the effect of air preheat temperature on flame characteristics in normal and microgravity conditions. We have obtained qualitative (global flame features) and some quantitative information on the features of flames using high temperature combustion air under normal gravity conditions with propane and methane as the fuels. This data will be compared with the data under microgravity conditions. The specific focus under normal gravity conditions has been on determining the global flame features as well as the spatial distribution of OH, CH, and C2 from flames using high temperature combustion air at different equivalence ratio.

  8. Gaseous and Particulate Matter Emissions of a Supercharged Spark Ignited Hydrogen Fueled Internal Combustion Engine

    NASA Astrophysics Data System (ADS)

    Kieran, Sean

    A spark ignited hydrogen fueled engine was operated at three equivalence ratios (0.4, 0.5, and 0.6) with a supercharger. During steady-state road load conditions, the engine produced exceptionally low unburned hydrocarbon, carbon monoxide, carbon dioxide, and particulate matter emissions. The oxides of nitrogen (NOx) emissions of the supercharged engine were 31.4, 149.5, and 787.0 mg*NOx/km for the equivalence ratios 0.4, 0.5, and 0.6 respectively. Given that the current EPA regulations are 99.4 mg*NOx/km, this engine configuration represents a possible replacement option for gasoline fueled engines without the need for exhaust after treatment. During engine start-up, some of the supercharged tests exhibited particulate matter emission spikes. These particulate matter spikes do not seem to be related to equivalence ratio, coolant temperature, testing order, or start-up acceleration. Currently, there is no explanation why some of the tests produced particulate matter during engine start-up and others did not.

  9. Thrust-vector control of a three-axis stabilized upper-stage rocket with fuel slosh dynamics

    NASA Astrophysics Data System (ADS)

    Rubio Hervas, Jaime; Reyhanoglu, Mahmut

    2014-05-01

    This paper studies the thrust vector control problem for an upper-stage rocket with fuel slosh dynamics. The dynamics of a three-axis stabilized spacecraft with a single partially-filled fuel tank are formulated and the sloshing propellant is modeled as a multi-mass-spring system, where the oscillation frequencies of the mass-spring elements represent the prominent sloshing modes. The equations of motion are expressed in terms of the three-dimensional spacecraft translational velocity vector, the attitude, the angular velocity, and the internal coordinates representing the slosh modes. A Lyapunov-based nonlinear feedback control law is proposed to control the translational velocity vector and the attitude of the spacecraft, while attenuating the sloshing modes characterizing the internal dynamics. A simulation example is included to illustrate the effectiveness of the control law.

  10. Designing the SSTO rocket

    NASA Astrophysics Data System (ADS)

    Payton, Gary; Sponable, Jess M.

    1991-04-01

    A review is presented of single-stage-to-orbit (SSTO) rocket vehicle structures fabricated with off-the-shelf technology. The parallel development of the Advanced Launch Development Program (ALDP) and NASP technology gave many spinoffs required for supportable, robust SSTO rockets. The advanced materials, structures, subsystems, cryogenic tanks, rockets, CFDs, and design tools are well suited to SSTO rockets. Developments include lightweight thermoplastic tanks for hydrogen fuel, lightweight carbon-carbon panels able to withstand 3,000 F reentry temperatures, and the CFD analysis of linear rocket thrusters installed on a ballistic rocket. The ALDP has also provided operations concepts, manufacturing technologies, and materials such as aluminum-lithium.

  11. Computer model predictions of the local effects of large, solid-fuel rocket motors on stratospheric ozone. Technical report

    SciTech Connect

    Zittel, P.F.

    1994-09-10

    The solid-fuel rocket motors of large space launch vehicles release gases and particles that may significantly affect stratospheric ozone densities along the vehicle's path. In this study, standard rocket nozzle and flowfield computer codes have been used to characterize the exhaust gases and particles through the afterburning region of the solid-fuel motors of the Titan IV launch vehicle. The models predict that a large fraction of the HCl gas exhausted by the motors is converted to Cl and Cl2 in the plume afterburning region. Estimates of the subsequent chemistry suggest that on expansion into the ambient daytime stratosphere, the highly reactive chlorine may significantly deplete ozone in a cylinder around the vehicle track that ranges from 1 to 5 km in diameter over the altitude range of 15 to 40 km. The initial ozone depletion is estimated to occur on a time scale of less than 1 hour. After the initial effects, the dominant chemistry of the problem changes, and new models are needed to follow the further expansion, or closure, of the ozone hole on a longer time scale.

  12. The second X-15 rocket plane (56-6671) is shown with two external fuel tanks which were added during

    NASA Technical Reports Server (NTRS)

    1965-01-01

    The second X-15 rocket plane (56-6671) is shown with two external fuel tanks which were added during its conversion to the X-15A-2 configuration in the mid-1960's. After receiving an ablative coating to protect the craft from the high temperatures associated with high-Mach-number supersonic flight, the X-15A-2 was then covered with a white sealant coat. This ablative coating and sealant and the additional fuel would help Air Force Col. William J. 'Pete' Knight fly the #2 X-15 to a world record speed of 4,520 mph (Mach 6.7). The famed X-15 rocket planes were flown at NASA's Dryden Flight Research Center, Edwards, California, from 1959 through 1968. The X-15 was developed to provide data on aerodynamics, structures, flight controls and the physiological aspects of high speed, high altitude flight. The joint NASA/U.S. Air Force/North American Aviation X-15 hypersonic flight research program is still considered to be one of the most successful NASA aeronautical research programs ever flown. One of the most advanced aeronautical tools of its day, the X-15 carried almost 600 instruments and sensors to record flight data. A wide range of experiments flown by the three X-15s during the 199-flight program helped advance the development of vital aeronautic and space flight systems.

  13. Gaseous fuel production from nonrecyclable paper wastes by using supported metal catalysts in high-temperature liquid water.

    PubMed

    Yamaguchi, Aritomo; Hiyoshi, Norihito; Sato, Osamu; Bando, Kyoko K; Shirai, Masayuki

    2010-06-21

    Paper wastes are used for the production of gaseous fuels over supported metal catalysts. The gasification of the nonrecyclable paper wastes, such as shredded documents and paper sludge, is carried out in high-temperature liquid water. The order of the catalytic activity for the gasification is found to be ruthenium>rhodium>platinum>palladium. A charcoal-supported ruthenium catalyst (Ru/C) is the most effective for the gasification of paper and cellulose. Paper wastes are gasified to a limited degree (32.6 carbon %) for 30 min in water at 523 K to produce methane and carbon dioxide, with a small amount of hydrogen. At 573 K, more complete gasification with almost 100 carbon % is achieved within 10 min in water. At 523 K, the gas yield of paper gasification over Ru/C is higher than that of cellulose powder. The gas yields are increased by ball-milling treatment of the recycled paper and cellulose powder. Printed paper wastes are also gasified at 523 K in water.

  14. Bonded and Sealed External Insulations for Liquid-Hydrogen-Fueled Rocket Tanks During Atmospheric Flight

    NASA Technical Reports Server (NTRS)

    Gray, V. H.; Gelder, T. F.; Cochran, R. P.; Goodykoontz, J. H.

    1960-01-01

    Several currently available nonmetallic insulation materials that may be bonded onto liquid-hydrogen tanks and sealed against air penetration into the insulation have been investigated for application to rockets and spacecraft. Experimental data were obtained on the thermal conductivities of various materials in the cryogenic temperature range, as well as on the structural integrity and ablation characteristics of these materials at high temperatures occasioned by aerodynamic heating during atmospheric escape. Of the materials tested, commercial corkboard has the best overall properties for the specific requirements imposed during atmospheric flight of a high-acceleration rocket vehicle.

  15. Done in 60 seconds- See a Massive Rocket Fuel Tank Built in A Minute

    NASA Image and Video Library

    2016-08-18

    The 7.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the rocket.

  16. Investigation of Non-Conventional Bio-Derived Fuels for Hybrid Rocket Motors

    DTIC Science & Technology

    2007-08-01

    estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the...data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this...1 I.II Review of Outside Hybrid Rocket Research

  17. Limits of thermochemical and photochemical syntheses of gaseous fuels: a finite-time thermodynamic analysis. Annual report, January-December 1985

    SciTech Connect

    Berry, R.S.

    1986-01-28

    The methods of evaluating processes for gaseous-fuel synthesis are generalized and applied to complex thermochemical processes, to photochemical and biophotochemical processes, and to electrolytic processes. Thermochemical methods do not appear promising in thermodynamic terms, even with careful strategies of heat transfer. Photochemical processes may be more advantageous, and electrochemical processes best of all, if no significant costs are associated with the production of electricity. The part of the project devoted to optimization of paths deals with finding optimal currents, as functions of time, for various fuel cells.

  18. Fuel Chemistry And Combustion Distribution Effects On Rocket Engine Combustion Stability

    DTIC Science & Technology

    2015-11-19

    experiments of Crowe et al. [50] are performed with PBAN- based polymer type solid propellants containing 16% aluminum particles. Crowe et al.’s nozzle...controlled by the combustion process at the propellant surface. As described above, the mass flow ratio and momentum ratio are calculated based on the...model rocket combustor, and associated modeling of particles in combustion chamber gas flows. 15. SUBJECT TERMS energetic propellants , combustion

  19. Econometric comparisons of liquid rocket engines for dual-fuel advanced earth-to-orbit shuttles

    NASA Technical Reports Server (NTRS)

    Martin, J. A.

    1978-01-01

    Econometric analyses of advanced Earth-to-orbit vehicles indicate that there are economic benefits from development of new vehicles beyond the space shuttle as traffic increases. Vehicle studies indicate the advantage of the dual-fuel propulsion in single-stage vehicles. This paper shows the economic effect of incorporating dual-fuel propulsion in advanced vehicles. Several dual-fuel propulsion systems are compared to a baseline hydrogen and oxygen system.

  20. Econometric comparisons of liquid rocket engines for dual-fuel advanced earth-to-orbit shuttles

    NASA Technical Reports Server (NTRS)

    Martin, J. A.

    1978-01-01

    Econometric analyses of advanced Earth-to-orbit vehicles indicate that there are economic benefits from development of new vehicles beyond the space shuttle as traffic increases. Vehicle studies indicate the advantage of the dual-fuel propulsion in single-stage vehicles. This paper shows the economic effect of incorporating dual-fuel propulsion in advanced vehicles. Several dual-fuel propulsion systems are compared to a baseline hydrogen and oxygen system.

  1. Rocket pollution reduction system

    SciTech Connect

    Geisler, R.L.

    1994-01-04

    A system is provided for reducing the emissions of hydrochloric acid (HCl) from solid fuel rockets, especially during ground disposal. An aqueous solution of an alkali metal hydroxide is injected as a mist into the rocket chamber as the rocket fuel is burned. The reaction of the alkali metal with hydrogen chloride (HCl) produces a salt and thereby minimizes the presence of hydrochloric acid in the rocket exhaust. An injected neutralizing material which reduces hydrochloric acid, but which produces less thrust than an equal weight of rocket fuel, can be injected into an operating rocket which carries a payload high above the earth, with the injected material being injected only while the rocket is at a lower altitude when hydrochloric acid is most undesirable. The injected material can be produced by a small auxiliary rocket device whose exhaust is delivered directly to the main rocket chamber, and with the exhaust of the auxiliary rocket device including a high proportion of magnesium to react with the hydrochloric acid with minimal degradation of rocket performance. 4 figs.

  2. Nuclear Rocket Ceramic Metal Fuel Fabrication Using Tungsten Powder Coating and Spark Plasma Sintering

    NASA Technical Reports Server (NTRS)

    Barnes, M. W.; Tucker, D. S.; Hone, L.; Cook, S.

    2017-01-01

    Nuclear thermal propulsion is an enabling technology for crewed Mars missions. An investigation was conducted to evaluate spark plasma sintering (SPS) as a method to produce tungsten-depleted uranium dioxide (W-dUO2) fuel material when employing fuel particles that were tungsten powder coated. Ceramic metal fuel wafers were produced from a blend of W-60vol% dUO2 powder that was sintered via SPS. The maximum sintering temperatures were varied from 1,600 to 1,850 C while applying a 50-MPa axial load. Wafers exhibited high density (>95% of theoretical) and a uniform microstructure (fuel particles uniformly dispersed throughout tungsten matrix).

  3. A simplified model for hybrid rocket performance prediction

    NASA Astrophysics Data System (ADS)

    Wolf, Robert S.; Wagner, John W.

    1992-02-01

    A computer code to predict hybrid rocket performance was developed and validated. The algorithm used is a simplification of the model derived by Marxman and Wooldridge. This model assumes the fuel regression rate to be controlled by convective heat transfer to the solid fuel from a relatively thin diffusion flame in a turbulent boundary layer. The model further assumes that the Reynolds analogy applies with mass addition at the wall. The computer code incorporates variable combustion product properties (temperature, molecular weight, and ratio of specific heats) as a function of the instantaneous global oxidizer/fuel ratio. The code was validated by constructing and firing a hybrid rocket motor. This motor used gaseous oxygen and hydroxyl-terminated polybutadiene as propellants. The oxygen flow rate used in the test was given as an input to the computer code, which then calculated chamber pressures and thrust. The agreement between test data and computer predictions was excellent.

  4. High Frequency Combustion Instability Studies of LOX/Methane Fueled Rocket Engines

    DTIC Science & Technology

    2009-09-25

    drops, secondary atomization of drops, drop heating and vaporization , mixing processes involving the drops and gases, mixture ratio distribution in space... hydrazine family of fuels. Engines using LOX/H2 propellants have experienced considerably less incidences of combustion instability, but are not totally...fuel tanks and its low vaporization temperature limiting its storability. Recently, renewed interest in lower operational costs, higher propellant

  5. 71. VIEW OF FUEL APRON FROM THE NORTHWEST. LEFT TO ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    71. VIEW OF FUEL APRON FROM THE NORTHWEST. LEFT TO RIGHT: HELIUM TANKS, GASEOUS NITROGEN TANKS, DIESEL FUEL TANK AND BACKUP GENERATOR, AND ROCKET FUEL TANKS. NORTHWEST CORNER OF THE LSB (BLDG. 751) AND LAUNCHER IN BACKGROUND ON LEFT; SOUTH CAMERA TOWER IN BACKGROUND ON RIGHT. - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 East, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

  6. Liquid rocket combustor computer code development

    NASA Technical Reports Server (NTRS)

    Liang, P. Y.

    1985-01-01

    The Advanced Rocket Injector/Combustor Code (ARICC) that has been developed to model the complete chemical/fluid/thermal processes occurring inside rocket combustion chambers are highlighted. The code, derived from the CONCHAS-SPRAY code originally developed at Los Alamos National Laboratory incorporates powerful features such as the ability to model complex injector combustion chamber geometries, Lagrangian tracking of droplets, full chemical equilibrium and kinetic reactions for multiple species, a fractional volume of fluid (VOF) description of liquid jet injection in addition to the gaseous phase fluid dynamics, and turbulent mass, energy, and momentum transport. Atomization and droplet dynamic models from earlier generation codes are transplated into the present code. Currently, ARICC is specialized for liquid oxygen/hydrogen propellants, although other fuel/oxidizer pairs can be easily substituted.

  7. Register of specialized sources for information on selected fuels and oxidizers. [rocket propellants, bibliographies

    NASA Technical Reports Server (NTRS)

    Ludtke, P. R.

    1975-01-01

    Thirty-eight (38) organizations are listed and described that catalog and file information in their data systems on fuel and oxidizers. The fuels include hydrogen, methane and hydrazine-type fuels; the oxidizers include oxygen, fluorine, flox, nitrogen tetroxide and ozone. The type of available information covers thermophysical properties, propellant systems, propellant fires-control-extinguishment, propellant explosions, propellant combustion, propellant safety, and fluorine chemistry. These organizations have assembled and collated their information so that it will be useful in the solution of engineering problems.

  8. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    NASA Technical Reports Server (NTRS)

    Emrich, William J., Jr.

    2008-01-01

    To support the eventual development of a nuclear thermal rocket engine, a state-of-the-art experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The test device simulates the environmental conditions (minus the radiation) to which nuclear rocket fuel components will be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner as to accurately reproduce the temperatures and heat fluxes normally expected to occur as a result of nuclear fission while at the same time being exposed to flowing hydrogen. This project is referred to as the Nuclear Thermal Rocket Element Environment Simulator or NTREES. The NTREES device is located at the Marshall Space flight Center in a laboratory which has been modified to accommodate the high powers required to heat the test articles to the required temperatures and to handle the gaseous hydrogen flow required for the tests. Other modifications to the laboratory include the installation of a nitrogen gas supply system and a cooling water supply system. During the design and construction of the facility, every effort was made to comply with all pertinent regulations to provide assurance that the facility could be operated in a safe and efficient manner. The NTREES system can currently supply up to 50 kW of inductive heating to the fuel test articles, although the facility has been sized to eventually allow test article heating levels of up to several megawatts.

  9. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    NASA Astrophysics Data System (ADS)

    Emrich, William J.

    2008-01-01

    To support a potential future development of a nuclear thermal rocket engine, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The test device simulates the environmental conditions (minus the radiation) to which nuclear rocket fuel components could be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner as to accurately reproduce the temperatures and heat fluxes normally expected to occur as a result of nuclear fission while at the same time being exposed to flowing hydrogen. This project is referred to as the Nuclear Thermal Rocket Element Environment Simulator or NTREES. The NTREES device is located at the Marshall Space flight Center in a laboratory which has been modified to accommodate the high powers required to heat the test articles to the required temperatures and to handle the gaseous hydrogen flow required for the tests. Other modifications to the laboratory include the installation of a nitrogen gas supply system and a cooling water supply system. During the design and construction of the facility, every effort was made to comply with all pertinent regulations to provide assurance that the facility could be operated in a safe and efficient manner. The NTREES system can currently supply up to 50 kW of inductive heating to the fuel test articles, although the facility has been sized to eventually allow test article heating levels of up to several megawatts.

  10. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    SciTech Connect

    Emrich, William J. Jr.

    2008-01-21

    To support a potential future development of a nuclear thermal rocket engine, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The test device simulates the environmental conditions (minus the radiation) to which nuclear rocket fuel components could be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner as to accurately reproduce the temperatures and heat fluxes normally expected to occur as a result of nuclear fission while at the same time being exposed to flowing hydrogen. This project is referred to as the Nuclear Thermal Rocket Element Environment Simulator or NTREES. The NTREES device is located at the Marshall Space flight Center in a laboratory which has been modified to accommodate the high powers required to heat the test articles to the required temperatures and to handle the gaseous hydrogen flow required for the tests. Other modifications to the laboratory include the installation of a nitrogen gas supply system and a cooling water supply system. During the design and construction of the facility, every effort was made to comply with all pertinent regulations to provide assurance that the facility could be operated in a safe and efficient manner. The NTREES system can currently supply up to 50 kW of inductive heating to the fuel test articles, although the facility has been sized to eventually allow test article heating levels of up to several megawatts.

  11. Fuel containment and stability in the gas core nuclear rocket. Final report, April 15, 1993--April 14, 1994

    SciTech Connect

    Kammash, T.

    1996-02-01

    One of the most promising approaches to advanced propulsion that could meet the objectives of the Space Exploration Initiative (SEI) is the open cycle gas core nuclear rocket (GCR). The energy in this device is generated by a fissioning uranium plasma which heats, through radiation, a propellant that flows around the core and exits through a nozzle, thereby converting thermal energy into thrust. Although such a scheme can produce very attractive propulsion parameters in the form of high specific impulse and high thrust, it does suffer from serious physics and engineering problems that must be addressed if it is to become a viable propulsion system. Among the major problems that must be solved are the confinement of the uranium plasma, potential instabilities and control problems associated with the dynamics of the uranium core, and the question of startup and fueling of such a reactor. In this paper, the authors focus their attention on the problems of equilibria and stability of the uranium care, and examine the potential use of an externally applied magnetic field for these purposes. They find that steady state operation of the reactor is possible only for certain care profiles that may not be compatible with the radiative aspect of the system. The authors also find that the system is susceptible to hydrodynamic and acoustic instabilities that could deplete the uranium fuel in a short time if not properly suppressed.

  12. Fuel efficient hydrodynamic containment for gas core fission reactor rocket propulsion. Final report, September 30, 1992--May 31, 1995

    SciTech Connect

    Sforza, P.M.; Cresci, R.J.

    1997-05-31

    Gas core reactors can form the basis for advanced nuclear thermal propulsion (NTP) systems capable of providing specific impulse levels of more than 2,000 sec., but containment of the hot uranium plasma is a major problem. The initial phase of an experimental study of hydrodynamic confinement of the fuel cloud in a gas core fission reactor by means of an innovative application of a base injection stabilized recirculation bubble is presented. The development of the experimental facility, a simulated thrust chamber approximately 0.4 m in diameter and 1 m long, is described. The flow rate of propellant simulant (air) can be varied up to about 2 kg/sec and that of fuel simulant (air, air-sulfur hexafluoride) up to about 0.2 kg/sec. This scale leads to chamber Reynolds numbers on the same order of magnitude as those anticipated in a full-scale nuclear rocket engine. The experimental program introduced here is focused on determining the size, geometry, and stability of the recirculation region as a function of the bleed ratio, i.e. the ratio of the injected mass flux to the free stream mass flux. A concurrent CFD study is being carried out to aid in demonstrating that the proposed technique is practical.

  13. Conceptual design for a kerosene fuel-rich gas-generator of a turbopump-fed liquid rocket engine

    NASA Astrophysics Data System (ADS)

    Son, Min; Koo, Jaye; Cho, Won Kook; Lee, Eun Seok

    2012-10-01

    A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.

  14. Review of the potential of silanes as rocket/scramjet fuels

    NASA Astrophysics Data System (ADS)

    Hidding, Bernhard; Pfitzner, Michael; Simone, Domenico; Bruno, Claudio

    2008-07-01

    Experimental use as well as theoretical considerations regarding silanes as fuels for spacecrafts and supersonic flight are reviewed. The historical circumstances leading to the utilization of monosilane as a fuel additive for scramjets are highlighted and milestones in the chemical research on silanes are summarized. Recent developments such as the use of monosilane as an ignition aid in the NASA X-43A scramjet flights as well as general progress in silicon hydride research, including liquid higher silanes and the resulting potential for the propulsion field are discussed.

  15. Early Rockets

    NASA Image and Video Library

    1958-01-31

    This illustration shows the main characteristics of the Jupiter C launch vehicle and its payload, the Explorer I satellite. The Jupiter C, America's first successful space vehicle, launched the free world's first scientific satellite, Explorer 1, on January 31, 1958. The four-stage Jupiter C measured almost 69 feet in length. The first stage was a modified liquid fueled Redstone missile. This main stage was about 57 feet in length and 70 inches in diameter. Fifteen scaled down SERGENT solid propellant motors were used in the upper stages. A "tub" configuration mounted on top of the modified Redstone held the second and third stages. The second stage consisted of 11 rockets placed in a ring formation within the tub. Inserted into the ring of second stage rockets was a cluster of 3 rockets making up the third stage. A fourth stage single rocket and the satellite were mounted atop the third stage. This "tub", all upper stages, and the satellite were set spirning prior to launching. The complete upper assembly measured 12.5 feet in length. The Explorer I carried the radiation detection experiment designed by Dr. James Van Allen and discovered the Van Allen Radiation Belt.

  16. Parametric studies with an atmospheric diffusion model that assesses toxic fuel hazards due to the ground clouds generated by rocket launches

    NASA Technical Reports Server (NTRS)

    Stewart, R. B.; Grose, W. L.

    1975-01-01

    Parametric studies were made with a multilayer atmospheric diffusion model to place quantitative limits on the uncertainty of predicting ground-level toxic rocket-fuel concentrations. Exhaust distributions in the ground cloud, cloud stabilized geometry, atmospheric coefficients, the effects of exhaust plume afterburning of carbon monoxide CO, assumed surface mixing-layer division in the model, and model sensitivity to different meteorological regimes were studied. Large-scale differences in ground-level predictions are quantitatively described. Cloud alongwind growth for several meteorological conditions is shown to be in error because of incorrect application of previous diffusion theory. In addition, rocket-plume calculations indicate that almost all of the rocket-motor carbon monoxide is afterburned to carbon dioxide CO2, thus reducing toxic hazards due to CO. The afterburning is also shown to have a significant effect on cloud stabilization height and on ground-level concentrations of exhaust products.

  17. Rockets -- Part II.

    ERIC Educational Resources Information Center

    Leitner, Alfred

    1982-01-01

    If two rockets are identical except that one engine burns in one-tenth the time of the other (total impulse and initial fuel mass of the two engines being the same), which rocket will rise higher? Why? The answer to this question (part 1 response in v20 n6, p410, Sep 1982) is provided. (Author/JN)

  18. Rockets -- Part II.

    ERIC Educational Resources Information Center

    Leitner, Alfred

    1982-01-01

    If two rockets are identical except that one engine burns in one-tenth the time of the other (total impulse and initial fuel mass of the two engines being the same), which rocket will rise higher? Why? The answer to this question (part 1 response in v20 n6, p410, Sep 1982) is provided. (Author/JN)

  19. Deposit formation in hydrocarbon rocket fuels with an evaluation of a propane heat transfer correlation

    NASA Technical Reports Server (NTRS)

    Masters, P. A.; Aukerman, C. A.

    1982-01-01

    A high pressure fuel coking testing apparatus was designed and developed and was used to evaluate thermal decomposition limits and carbon decomposition rates in heated copper tubes for hydrocarbon fuels. A commercial propane (90% grade) and chemically pure (CP) propane were tested. Heat transfer to supercritical propane was evaluated at 136 atm, bulk fluid velocities of 6 to 30 m/s, and tube wall temperatures in the range of 422 to 811 K. A forced convection heat transfer correlation developed in a previous test effort verified a prediction of most of the experimental data within a + or - 30% range, with good agreement for the CP propane data. No significant differences were apparent in the predictions derived from the correlation when the carbon resistance was included with the film resistance. A post-test scanning electron microprobe analysis indicated occurrences of migration and interdiffusion of copper into the carbon deposit.

  20. System Sensitivity Analysis Applied to the Conceptual Design of a Dual-Fuel Rocket SSTO

    NASA Technical Reports Server (NTRS)

    Olds, John R.

    1994-01-01

    This paper reports the results of initial efforts to apply the System Sensitivity Analysis (SSA) optimization method to the conceptual design of a single-stage-to-orbit (SSTO) launch vehicle. SSA is an efficient, calculus-based MDO technique for generating sensitivity derivatives in a highly multidisciplinary design environment. The method has been successfully applied to conceptual aircraft design and has been proven to have advantages over traditional direct optimization methods. The method is applied to the optimization of an advanced, piloted SSTO design similar to vehicles currently being analyzed by NASA as possible replacements for the Space Shuttle. Powered by a derivative of the Russian RD-701 rocket engine, the vehicle employs a combination of hydrocarbon, hydrogen, and oxygen propellants. Three primary disciplines are included in the design - propulsion, performance, and weights & sizing. A complete, converged vehicle analysis depends on the use of three standalone conceptual analysis computer codes. Efforts to minimize vehicle dry (empty) weight are reported in this paper. The problem consists of six system-level design variables and one system-level constraint. Using SSA in a 'manual' fashion to generate gradient information, six system-level iterations were performed from each of two different starting points. The results showed a good pattern of convergence for both starting points. A discussion of the advantages and disadvantages of the method, possible areas of improvement, and future work is included.

  1. Rocket Propellant Ducts (Cryogenic Fuel Lines): First Cut Approximations and Design Guidance

    NASA Technical Reports Server (NTRS)

    Brewer, William V.

    1998-01-01

    The design team has to set parameters before analysis can take place. Analysis is customarily a thorough and time consuming process which can take weeks or even months. Only when analysis is complete can the designer obtain feedback. If margins are negative, the process must be repeated to a greater or lesser degree until satisfactory results are achieved. Reduction of the number of iterations thru this loop would beneficially conserve time and resources. The task was to develop relatively simple, easy to use, guidelines and analytic tools that allow the designer to evaluate what effect various alternatives may have on performance as the design progresses. "Easy to use" is taken to mean closed form approximations and the use of graphic methods. "Simple" implies that 2-d and quasi 3-d approximations be exploited to whatever degree is useful before more resource intensive methods are applied. The objective is to avoid the grosser violation of performance margins at the outset. Initial efforts are focused on thermal expansion/contraction and rigid body kinematics as they relate to propellant duct displacements in the gimbal plane loop (GPL). The purpose of the loop is to place two flexible joints on the same two orthogonal intersecting axes as those of the rocket motor gimbals. This supposes the ducting will flex predictably with independent rotations corresponding to those of the motor gimbal actions. It can be shown that if GPL joint axes do not coincide with motor gimbal axes, displacement incompatibilities result in less predictable movement of the ducts.

  2. Investigation of a Tricarbide Grooved Ring Fuel Element for a Nuclear Thermal Rocket

    NASA Technical Reports Server (NTRS)

    Taylor, Brian D.; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

    2017-01-01

    Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the temperature limitations of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment.

  3. [IMMUNOCYTOCHEMICAL ANALYSIS OF THE DISTURBANCES IN THE STRUCTURE OF SYNAPTONEMAL COMPLEXES IN SPERMATOCYTE NUCLEI IN MICE UNDER EXPOSURE TO ROCKET FUEL COMPONENT].

    PubMed

    Lovinskaya, A V; Kolumbayeva, S Zh; Abilev, S K; Kolomiets, O L

    2016-01-01

    There was performed an assessment of genotoxic effects of rocket fuel component--unsymmetrical dimethylhydrazine (UDMH, heptyl)--on forming germ cells of male mice. Immunocytochemically there was studied the structure of meiotic nuclei at different times after the intraperitoneal administration of UDMH to male mice. There were revealed following types of disturbances of the structure of synaptonemal complexes (SCs) of meiotic chromosomes: single and multiple fragments of SCs associations of autosomes with a sex bivalent, atypical structure of the SCs with a frequency higher than the reference level. In addition, there were found the premature desinapsis of sex bivalents, the disorder offormation of the genital corpuscle and ring SCs. Established disorders in SCs of spermatocytes, analyzed at 38th day after the 10-days intoxication of animal by the component of rocket fuel, attest to the risk of permanent persistence of chromosomal abnormalities occurring in the pool of stem cells.

  4. Mixing characteristics of injector elements in liquid rocket engines - A computational study

    NASA Astrophysics Data System (ADS)

    Lohr, Jonathan C.; Trinh, Huu P.

    1992-07-01

    A computational study has been performed to better understand the mixing characteristics of liquid rocket injector elements. Variations in injector geometry as well as differences in injector element inlet flow conditions are among the areas examined in the study. Most results involve the nonreactive mixing of gaseous fuel with gaseous oxidizer but preliminary results are included that involve the spray combustion of oxidizer droplets. The purpose of the study is to numerically predict flowfield behavior in individual injector elements to a high degree of accuracy and in doing so to determine how various injector element properties affect the flow.

  5. Expected Detection Limits of Hydrazine-Based Rocket Fuels and Their Selected Oxidation Products by 12C16O2 Laser Spectroscopic Techniques.

    DTIC Science & Technology

    1980-08-15

    DAO089 b AEROSPACE CORP EL SEGUNDO CA CHEMISTRY AND PHYSICS LAB F/6 21/9.1 ExpECTED DETECTION LIMITS OF HYDRAZINE-BASED ROCKET FUELS AND T--ETC(UW...n ( aV Pin r I Pi n + i, n- Il On ) Pn n mnr The m simultaneous equations can be compacted to a set of n simultaneous equations so all m of

  6. Reynolds-averaged Navier-Stokes analysis of the flow through a model rocket-based combined-cycle engine with an independently-fueled ramjet stream

    NASA Astrophysics Data System (ADS)

    Bond, Ryan Bomar

    A new concept for the low speed propulsion mode in rocket based combined cycle (RBCC) engines has been developed as part of the NASA GTX program. This concept, called the independent ramjet stream (IRS) cycle, is a variation of the traditional ejector ramjet (ER) design and involves the injection of hydrogen fuel directly into the air stream, where it is ignited by the rocket plume. Experiments and computational fluid dynamics (CFD) are currently being used to evaluate the feasibility of the new design. In this work, a Navier-Stokes code valid for general reactive flows is applied to the model engine under cold flow, ejector ramjet, and IRS cycle operation. Pressure distributions corresponding to cold-flow and ejector ramjet operation are compared with experimental data. The engine response under independent ramjet stream cycle operation is examined for different reaction models and grid sizes. The engine response to variations in fuel injection is also examined. Mode transition simulations are also analyzed both with and without a nitrogen purge of the rocket. The solutions exhibit a high sensitivity to both grid resolution and reaction mechanism, but they do indicate that thermal throat ramjet operation is possible through the injection and burning of additional fuel into the air stream. The solutions also indicate that variations in fuel injection location can affect the position of the thermal throat. The numerical simulations predicted successful mode transition both with and without a nitrogen purge of the rocket; however, the reliability of the mode transition results cannot be established without experimental data to validate the reaction mechanism.

  7. Theoretical rocket performance of JP-4 fuel with mixtures of liquid ozone and fluorine

    NASA Technical Reports Server (NTRS)

    Huff, Vearl N; Gordon, Sanford

    1957-01-01

    Data were estimated by means of a heat correction equation using data for JP-4 fuel with mixtures of oxygen and flourine. The estimated data were checked for several cases by direct calculations. The difference in specific impulse between the estimated and directly calculated values was from 0.2 to 0.8 pound-second per pound. The maximum value of specific impulse was 334.9 pound-seconds per pound for a combustion-chamber pressure of 600 pounds per square inch absolute and an exit pressure of 1 atmosphere.

  8. Rocket Fuel R and D at AFRL: Recent Activities and Future Direction

    DTIC Science & Technology

    2017-04-12

    4, 540 klbf • GG • 190 klbf • Merlin CH4 RP • ORSC • RD-170/180, 800 klbf RP Meanwhile, in Russia… U.S. hasn’t developed or built high- pressure LOX/HC...1 10 100 1000 10000 100000 Te m pe ra tu re (F ) Pressure (psia) 20DISTRIBUTION A: Approved for Public Release; Distribution Unlimited PA...17163 Combustion Chemistry Models 1410 K, 2.3 atm, 457 ppm n-C12H26/O2/Ar, φ = 1 Ignition ~20 μs ~700 μs Fuel pyrolysis Fragments oxidation Physics

  9. Further analytical study of hybrid rocket combustion

    NASA Technical Reports Server (NTRS)

    Hung, W. S. Y.; Chen, C. S.; Haviland, J. K.

    1972-01-01

    Analytical studies of the transient and steady-state combustion processes in a hybrid rocket system are discussed. The particular system chosen consists of a gaseous oxidizer flowing within a tube of solid fuel, resulting in a heterogeneous combustion. Finite rate chemical kinetics with appropriate reaction mechanisms were incorporated in the model. A temperature dependent Arrhenius type fuel surface regression rate equation was chosen for the current study. The governing mathematical equations employed for the reacting gas phase and for the solid phase are the general, two-dimensional, time-dependent conservation equations in a cylindrical coordinate system. Keeping the simplifying assumptions to a minimum, these basic equations were programmed for numerical computation, using two implicit finite-difference schemes, the Lax-Wendroff scheme for the gas phase, and, the Crank-Nicolson scheme for the solid phase.

  10. ELM - A SIMPLE TOOL FOR THERMAL-HYDRAULIC ANALYSIS OF SOLID-CORE NUCLEAR ROCKET FUEL ELEMENTS

    NASA Technical Reports Server (NTRS)

    Walton, J. T.

    1994-01-01

    ELM is a simple computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in nuclear thermal rockets. Written for the nuclear propulsion project of the Space Exploration Initiative, ELM evaluates the various heat transfer coefficient and friction factor correlations available for turbulent pipe flow with heat addition. In the past, these correlations were found in different reactor analysis codes, but now comparisons are possible within one program. The logic of ELM is based on the one-dimensional conservation of energy in combination with Newton's Law of Cooling to determine the bulk flow temperature and the wall temperature across a control volume. Since the control volume is an incremental length of tube, the corresponding pressure drop is determined by application of the Law of Conservation of Momentum. The size, speed, and accuracy of ELM make it a simple tool for use in fuel element parametric studies. ELM is a machine independent program written in FORTRAN 77. It has been successfully compiled on an IBM PC compatible running MS-DOS using Lahey FORTRAN 77, a DEC VAX series computer running VMS, and a Sun4 series computer running SunOS UNIX. ELM requires 565K of RAM under SunOS 4.1, 360K of RAM under VMS 5.4, and 406K of RAM under MS-DOS. Because this program is machine independent, no executable is provided on the distribution media. The standard distribution medium for ELM is one 5.25 inch 360K MS-DOS format diskette. ELM was developed in 1991. DEC, VAX, and VMS are trademarks of Digital Equipment Corporation. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation.

  11. ELM - A SIMPLE TOOL FOR THERMAL-HYDRAULIC ANALYSIS OF SOLID-CORE NUCLEAR ROCKET FUEL ELEMENTS

    NASA Technical Reports Server (NTRS)

    Walton, J. T.

    1994-01-01

    ELM is a simple computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in nuclear thermal rockets. Written for the nuclear propulsion project of the Space Exploration Initiative, ELM evaluates the various heat transfer coefficient and friction factor correlations available for turbulent pipe flow with heat addition. In the past, these correlations were found in different reactor analysis codes, but now comparisons are possible within one program. The logic of ELM is based on the one-dimensional conservation of energy in combination with Newton's Law of Cooling to determine the bulk flow temperature and the wall temperature across a control volume. Since the control volume is an incremental length of tube, the corresponding pressure drop is determined by application of the Law of Conservation of Momentum. The size, speed, and accuracy of ELM make it a simple tool for use in fuel element parametric studies. ELM is a machine independent program written in FORTRAN 77. It has been successfully compiled on an IBM PC compatible running MS-DOS using Lahey FORTRAN 77, a DEC VAX series computer running VMS, and a Sun4 series computer running SunOS UNIX. ELM requires 565K of RAM under SunOS 4.1, 360K of RAM under VMS 5.4, and 406K of RAM under MS-DOS. Because this program is machine independent, no executable is provided on the distribution media. The standard distribution medium for ELM is one 5.25 inch 360K MS-DOS format diskette. ELM was developed in 1991. DEC, VAX, and VMS are trademarks of Digital Equipment Corporation. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation.

  12. Large eddy simulation of combustion characteristics in a kerosene fueled rocket-based combined-cycle engine combustor

    NASA Astrophysics Data System (ADS)

    Huang, Zhi-wei; He, Guo-qiang; Qin, Fei; Cao, Dong-gang; Wei, Xiang-geng; Shi, Lei

    2016-10-01

    This study reports combustion characteristics of a rocket-based combined-cycle engine combustor operating at ramjet mode numerically. Compressible large eddy simulation with liquid kerosene sprayed and vaporized is used to study the intrinsic unsteadiness of combustion in such a propulsion system. Results for the pressure oscillation amplitude and frequency in the combustor as well as the wall pressure distribution along the flow-path, are validated using experimental data, and they show acceptable agreement. Coupled with reduced chemical kinetics of kerosene, results are compared with the simultaneously obtained Reynolds-Averaged Navier-Stokes results, and show significant differences. A flow field analysis is also carried out for further study of the turbulent flame structures. Mixture fraction is used to determine the most probable flame location in the combustor at stoichiometric condition. Spatial distributions of the Takeno flame index, scalar dissipation rate, and heat release rate reveal that different combustion modes, such as premixed and non-premixed modes, coexisted at different sections of the combustor. The RBCC combustor is divided into different regions characterized by their non-uniform features. Flame stabilization mechanism, i.e., flame propagation or fuel auto-ignition, and their relative importance, is also determined at different regions in the combustor.

  13. Mechanical and Combustion Performance of Multi-Walled Carbon Nanotubes as an Additive to Paraffin-Based Solid Fuels for Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Larson, Daniel B.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth, K.; Koo, Joseph H.; Story, George

    2012-01-01

    Paraffin-based solid fuels for hybrid rocket motor applications are recognized as a fastburning alternative to other fuel binders such as HTPB, but efforts to further improve the burning rate and mechanical properties of paraffin are still necessary. One approach that is considered in this study is to use multi-walled carbon nanotubes (MWNT) as an additive to paraffin wax. Carbon nanotubes provide increased electrical and thermal conductivity to the solid-fuel grains to which they are added, which can improve the mass burning rate. Furthermore, the addition of ultra-fine aluminum particles to the paraffin/MWNT fuel grains can enhance regression rate of the solid fuel and the density impulse of the hybrid rocket. The multi-walled carbon nanotubes also present the possibility of greatly improving the mechanical properties (e.g., tensile strength) of the paraffin-based solid-fuel grains. For casting these solid-fuel grains, various percentages of MWNT and aluminum particles will be added to the paraffin wax. Previous work has been published about the dispersion and mixing of carbon nanotubes.1 Another manufacturing method has been used for mixing the MWNT with a phenolic resin for ablative applications, and the manufacturing and mixing processes are well-documented in the literature.2 The cost of MWNT is a small fraction of single-walled nanotubes. This is a scale-up advantage as future applications and projects will require low cost additives to maintain cost effectiveness. Testing of the solid-fuel grains will be conducted in several steps. Dog bone samples will be cast and prepared for tensile testing. The fuel samples will also be analyzed using thermogravimetric analysis and a high-resolution scanning electron microscope (SEM). The SEM will allow for examination of the solid fuel grain for uniformity and consistency. The paraffin-based fuel grains will also be tested using two hybrid rocket test motors located at the Pennsylvania State University s High Pressure

  14. Raman Gas Species Measurements in Hydrocarbon-Fueled Rocket Engine Injector Flows

    NASA Technical Reports Server (NTRS)

    Wehrmeyer, Joseph A.; Trinh, Huu Phuoc; Hartfield, Roy J.; Dobson, Christopher C.; Eskridge, Richard H.

    2000-01-01

    Propellent injector development at MSFC (Marshall Space Flight Center) includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellent mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

  15. Hydrocarbon-Fueled Rocket Plume Measurement Using Polarized UV Raman Spectroscopy

    NASA Technical Reports Server (NTRS)

    Wehrmeyer, Joseph A.

    2002-01-01

    The influence of pressure upon the signal strength and polarization properties of UV Raman signals has been investigated experimentally up to pressures of 165 psia (11 atm). No significant influence of pressure upon the Raman scattering cross section or depolarization ratio of the N2 Raman signal was found. The Raman scattering signal varied linearly with pressure for the 300 K N2 samples examined, thus showing no enhancement of cross section with increasing pressure. However at the highest pressures associated with rocket engine combustion, there could be an increase in the Raman scattering cross section, based upon others' previous work at higher pressures than those examined in this work. The influence of pressure upon thick fused silica windows, used in the NASA Modular Combustion Test Article, was also investigated. No change in the transmission characteristics of the windows occurred as the pressure difference across the windows increased from 0 psig up to 150 psig. A calibration was performed on the UV Raman system at Vanderbilt University, which is similar to the one at the NASA-Marshall Test Stand 115. The results of this calibration are described in the form of temperature-dependent functions, f(T)'s, that account for the increase in Raman scattering cross section with an increase in temperature and also account for the reduction in collected Raman signal if wavelength integration does not occur across the entire wavelength range of the Raman signal. These functions generally vary only by approximately 10% across their respective temperature ranges, except for the case Of CO2, where there is a factor of three difference in its f(T) from 300 K to 2500 K. However this trend for CO2 is consistent with the experimental work of others, and is expected based on the low characteristic vibrational temperature Of CO2. A time-averaged temperature measurement technique has been developed, using the same equipment as for the work mentioned above, that is based upon

  16. Propellant vaporization as a criterion for rocket-engine design : experimental effect of fuel temperature on liquid-oxygen - heptane performance

    NASA Technical Reports Server (NTRS)

    Heidmann, M F

    1957-01-01

    Characteristic exhaust velocity of a 200-pound-thrust rocket engine was evaluated for fuel temperatures of -90 degrees, and 200 degrees f with a spray formed by two impinging heptane jets reacting in a highly atomized oxygen atmosphere. Tests covered a range of mixture ratios and chamber lengths. The characteristic exhaust-velocity efficiency increased 2 percent for a 290 degree f increase in fuel temperature. This increase in performance can be compared with that obtained by increasing chamber length by about 1/2 inch. The result agrees with the fuel-temperature effect predicted from an analysis based on droplet evaporation theory. Mixture ratio markedly affected characteristic exhaust velocity efficiency, but total flow rate and fuel temperature did not.

  17. Investigation of Critical Burning of Fuel Droplets. [of liquid rocket propellant

    NASA Technical Reports Server (NTRS)

    Chanin, S. P.; Shearer, A. J.; Faeth, G. M.

    1976-01-01

    An earlier analysis for the combustion response of a liquid monopropellant strand (hydrazine) was extended to consider individual droplets and sprays. While small drops gave low or negative response, large droplets provided response near unity at low frequencies, with the response declining at frequencies greater than the characteristic liquid phase frequency. Temperature gradients in the liquid phase resulted in response peaks greater than unity. A second response peak was found for large drops which corresponded to gas phase transient effects. Spray response was generally reduced from the response of the largest injected droplet, however, even a small percentage of large droplets can yield appreciable response. An apparatus was designed and fabricated to allow observation of bipropellant fuel spray combustion at elevated pressures. A locally homogeneous model was developed to describe this combustion process which allows for high pressure phenomena associated with the thermodynamic critical point.

  18. Nitrato-Functionalized Task-Specific Ionic Liquids as Attractive Hypergolic Rocket Fuels.

    PubMed

    Wang, Yi; Huang, Shi; Zhang, Wenquan; Liu, Tianlin; Qi, Xiujuan; Zhang, Qinghua

    2017-09-12

    Hypergolic ionic liquids (HILs) as potential replacements for hydrazine derivatives have attracted increasing interest over the last decade. Previous studies on HILs have mostly concentrated on the anionic innovations of ionic liquids to shorten the ignition delay (ID) time, but little attention has been paid to cationic modifications and their structure-property relationships. In this work, we present a new strategy of cationic functionalization by introducing the energetic nitrato group into the cationic units of HILs. Interestingly, the introduction of oxygen-rich nitrato groups into the cationic structure significantly improved the combustion performance of HILs with larger flame diameters and duration times. The density-specific impulse (ρIsp ) of these novel HILs are all above 279.0 s g cm(-3) , much higher than that of UDMH (215.7 s g cm(-3) ). In addition, the densities of these HILs are in the range of 1.22-1.39 g cm(-3) , which is much higher than that of UDMH (0.79 g cm(-3) ), showing their higher loading capacity than hydrazine-derived fuels in a propellant tank. This promising strategy of introducing nitrato groups into the cationic structures has provided a new platform for developing high-performing HILs with improved combustion properties. © 2017 Wiley-VCH Verlag GmbH & Co. KGaA, Weinheim.

  19. Experimental investigations on pulse detonation rocket engine with various injectors and nozzles

    NASA Astrophysics Data System (ADS)

    Yan, Yu; Fan, Wei; Wang, Ke; Zhu, Xu-dong; Mu, Yang

    2011-07-01

    Pulse detonation engines (PDEs) may represent a revolutionary approach to propulsion. The engine of simple construction can be easily manufactured. The pulse detonation rocket engine (PDRE) used here are 30 mm in inner diameter and 860 mm in length. Liquid kerosene, gaseous oxygen and nitrogen were used as fuel, oxidizer and purge gas, respectively. Two-phase detonation generating is harder than gaseous detonation due to liquid fuel atomization and mixing of two-phase reactants. It is a difficult task for liquid fuel and gaseous oxidizer to mix and form uniformly distributed mixture in the entire long engine during filling process in a short time. Therefore the velocities of fuel and oxidizer must be well designed to achieve not only the requirement of filling the entire engine but also the requirement of liquid fuel atomization and reactants mixing. Four injectors were tested to improve the atomization of liquid fuel and mixing process of reactants for performance enhancement of PDRE. Injector with small fuel exit area and large gas exit area was found to be effective for liquid fuel atomization and reactants mixing process. The PDRE with injector B performed the best among all the injectors tested. Nozzles are critical components in improving the performance of PDRE. Four kinds of bell-shaped converging-diverging nozzles were also tested here in order to enhance the performance of PDRE. It was found that a nozzle with high contraction ratio and high expansion ratio generated the highest thrust augmentation of 27.3%.

  20. Evaluation of advanced combustion concepts for dry NO sub x suppression with coal-derived, gaseous fuels

    NASA Technical Reports Server (NTRS)

    Beebe, K. W.; Symonds, R. A.; Notardonato, J. J.

    1982-01-01

    The emissions performance of a rich lean combustor (developed for liquid fuels) was determined for combustion of simulated coal gases ranging in heating value from 167 to 244 Btu/scf (7.0 to 10.3 MJ/NCM). The 244 Btu/scf gas is typical of the product gas from an oxygen blown gasifier, while the 167 Btu/scf gas is similar to that from an air blown gasifier. NOx performance of the rich lean combustor did not meet program goals with the 244 Btu/scf gas because of high thermal NOx, similar to levels expected from conventional lean burning combustors. The NOx emissions are attributed to inadequate fuel air mixing in the rich stage resulting from the design of the large central fuel nozzle delivering 71% of the total gas flow. NOx yield from ammonia injected into the fuel gas decreased rapidly with increasing ammonia level, and is projected to be less than 10% at NH3 levels of 0.5% or higher. NOx generation from NH3 is significant at ammonia concentrations significantly less than 0.5%. These levels may occur depending on fuel gas cleanup system design. CO emissions, combustion efficiency, smoke and other operational performance parameters were satisfactory. A test was completed with a catalytic combustor concept with petroleum distillate fuel. Reactor stage NOx emissions were low (1.4g NOx/kg fuel). CO emissions and combustion efficiency were satisfactory. Airflow split instabilities occurred which eventually led to test termination.

  1. Manufacturing of 5.5 Meter Diameter Cryogenic Fuel Tank Domes for the NASA Ares I Rocket

    NASA Technical Reports Server (NTRS)

    Jones, Ronald E.; Carter, Robert W.

    2012-01-01

    The Ares I rocket is the first launch vehicle scheduled for manufacture under the National Aeronautic and Space Administration s (NASA s) Constellation program. A series of full-scale Ares I development articles have been constructed on the Robotic Weld Tool at the NASA George C. Marshall Space Flight Center in Huntsville, Alabama. The Robotic Weld Tool is a 100 ton, 7-axis, robotic manufacturing system capable of machining and friction stir welding large-scale space hardware. This presentation will focus on the friction stir welding of 5.5m diameter cryogenic fuel tank components; specifically, the liquid hydrogen forward dome (LH2 MDA) and the common bulkhead manufacturing development articles (CBMDA). The LH2 MDA was the first full-scale, flight-like Ares I hardware produced under the Constellation Program. It is a 5.5m diameter elliptical dome assembly consisting of eight gore panels, a y-ring stiffener and a manhole fitting. All components are made from aluminum-lithium alloy 2195. Conventional and self-reacting friction stir welding was used on this article. Manufacturing solutions will be discussed including the implementation of photogrammetry, an advanced metrology technique, as well as fixtureless welding. The LH2 MDA is the first known fully friction stir welded dome ever produced. The completion of four Common Bulkhead Manufacturing Development Articles (CBMDA) will also be highlighted. Each CBMDA consists of a 5.5m diameter spun-formed dome friction stir welded to a y-ring stiffener. The domes and y-rings are made of aluminum 2014 and 2219 respectively. An overview of CBMDA manufacturing processes and the effect of tooling on weld defect formation will be discussed.

  2. Robotic Manufacturing of 5.5 Meter Cryogenic Fuel Tank Dome Assemblies for the NASA Ares I Rocket

    NASA Technical Reports Server (NTRS)

    Jones, Ronald E.

    2012-01-01

    The Ares I rocket is the first launch vehicle scheduled for manufacture under the National Aeronautic and Space Administration's (NASA's) Constellation program. A series of full-scale Ares I development articles have been constructed on the Robotic Weld Tool at the NASA George C. Marshall Space Flight Center in Huntsville, Alabama. The Robotic Weld Tool is a 100 ton, 7-axis, robotic manufacturing system capable of machining and friction stir welding large-scale space hardware. This presentation will focus on the friction stir welding of 5.5m diameter cryogenic fuel tank components; specifically, the liquid hydrogen forward dome (LH2 MDA), the common bulkhead manufacturing development articles (CBMDA) and the thermal protection system demonstration dome (TPS Dome). The LH2 MDA was the first full-scale, flight-like Ares I hardware produced under the Constellation Program. It is a 5.5m diameter elliptical dome assembly consisting of eight gore panels, a y-ring stiffener and a manhole fitting. All components are made from aluminumlithium alloy 2195. Conventional and self-reacting friction stir welding was used on this article. An overview of the manufacturing processes will be discussed. The LH2 MDA is the first known fully friction stir welded dome ever produced. The completion of four Common Bulkhead Manufacturing Development Articles (CBMDA) and the TPS Dome will also be highlighted. Each CBMDA and the TPS Dome consists of a 5.5m diameter spun-formed dome friction stir welded to a y-ring stiffener. The domes and y-rings are made of aluminum 2014 and 2219 respectively. The TPS Dome has an additional aluminum alloy 2195 barrel section welded to the y-ring. Manufacturing solutions will be discussed including "fixtureless" welding with self reacting friction stir welding.

  3. Gaseous detonations

    SciTech Connect

    Nettleton, M.A.

    1987-01-01

    Focusing predominantly on safety problems in handling combustible gas or dust mixtures with air or oxygen, the book is a reference on gaseous detonations. Topics covered include: unidimensional models, structure of detonation fronts, and interaction of a detonation with confinement.

  4. Fundamental Phenomena on Fuel Decomposition and Boundary-Layer Combustion Precesses with Applications to Hybrid Rocket Motors. Part 1; Experimental Investigation

    NASA Technical Reports Server (NTRS)

    Kuo, Kenneth K.; Lu, Yeu-Cherng; Chiaverini, Martin J.; Johnson, David K.; Serin, Nadir; Risha, Grant A.; Merkle, Charles L.; Venkateswaran, Sankaran

    1996-01-01

    This final report summarizes the major findings on the subject of 'Fundamental Phenomena on Fuel Decomposition and Boundary-Layer Combustion Processes with Applications to Hybrid Rocket Motors', performed from 1 April 1994 to 30 June 1996. Both experimental results from Task 1 and theoretical/numerical results from Task 2 are reported here in two parts. Part 1 covers the experimental work performed and describes the test facility setup, data reduction techniques employed, and results of the test firings, including effects of operating conditions and fuel additives on solid fuel regression rate and thermal profiles of the condensed phase. Part 2 concerns the theoretical/numerical work. It covers physical modeling of the combustion processes including gas/surface coupling, and radiation effect on regression rate. The numerical solution of the flowfield structure and condensed phase regression behavior are presented. Experimental data from the test firings were used for numerical model validation.

  5. Rheological, optical, and ballistic investigations of paraffin-based fuels for hybrid rocket propulsion using a two-dimensional slab-burner

    NASA Astrophysics Data System (ADS)

    Kobald, M.; Toson, E.; Ciezki, H.; Schlechtriem, S.; di Betta, S.; Coppola, M.; DeLuca, L.

    2016-07-01

    This paper describes combined rheological, ballistic, and optical analyses performed on paraffin-based mixtures that can be used as high regression rate hybrid rocket fuels. Experimental activities have been done at the DLR Institute of Space Propulsion in Lampoldshausen and at SPLab of Politecnico di Milano [1]. Herein, the experiments that were performed at the DLR are described in detail. Viscosity, surface tension, and regression rate of the fuels have been determined. Furthermore, the combustion was evaluated by optical measurements. Data collected so far indicate an increasing regression rate for decreasing viscosity of the liquid paraffin which is in accordance with the current theories. Droplet entrainment, which is related to high regression rates, is only visible for the low-viscosity paraffin-based fuels.

  6. Impact of alternative fuels on emissions characteristics of a gas turbine engine - part 1: gaseous and particulate matter emissions.

    PubMed

    Lobo, Prem; Rye, Lucas; Williams, Paul I; Christie, Simon; Uryga-Bugajska, Ilona; Wilson, Christopher W; Hagen, Donald E; Whitefield, Philip D; Blakey, Simon; Coe, Hugh; Raper, David; Pourkashanian, Mohamed

    2012-10-02

    Growing concern over emissions from increased airport operations has resulted in a need to assess the impact of aviation related activities on local air quality in and around airports, and to develop strategies to mitigate these effects. One such strategy being investigated is the use of alternative fuels in aircraft engines and auxiliary power units (APUs) as a means to diversify fuel supplies and reduce emissions. This paper summarizes the results of a study to characterize the emissions of an APU, a small gas turbine engine, burning conventional Jet A-1, a fully synthetic jet fuel, and other alternative fuels with varying compositions. Gas phase emissions were measured at the engine exit plane while PM emissions were recorded at the exit plane as well as 10 m downstream of the engine. Five percent reduction in NO(x) emissions and 5-10% reduction in CO emissions were observed for the alternative fuels. Significant reductions in PM emissions at the engine exit plane were achieved with the alternative fuels. However, as the exhaust plume expanded and cooled, organic species were found to condense on the PM. This increase in organic PM elevated the PM mass but had little impact on PM number.

  7. Alternative Fuels

    EPA Pesticide Factsheets

    Alternative fuels include gaseous fuels such as hydrogen, natural gas, and propane; alcohols such as ethanol, methanol, and butanol; vegetable and waste-derived oils; and electricity. Overview of alternative fuels is here.

  8. A preliminary assessment of the feasibility of deriving liquid and gaseous fuels from grown and waste organics

    NASA Technical Reports Server (NTRS)

    Graham, R. W.; Reynolds, T. W.; Hsu, Y.-Y.

    1976-01-01

    An estimate is obtained of the yearly supply of organic material for conversion to fuels, the energy potential is evaluated, and the fermentation and pyrolysis conversion processes are discussed. An investigation is conducted of the estimated cost of fuel from organics and the conclusions of an overall evaluation are presented. It is found that climate, land availability and economics of agricultural production and marketing, food demand, fertilizer shortage, and water availability combine to cast doubts on the feasibility of producing grown organic matter for fuel, in competition with food, feed, or fiber. Less controversial is the utilization of agricultural, industrial, and domestic waste as a conversion feedstock. The evaluation of a demonstration size system is recommended.

  9. KUGEL: a thermal, hydraulic, fuel performance, and gaseous fission product release code for pebble bed reactor core analysis

    SciTech Connect

    Shamasundar, B.I.; Fehrenbach, M.E.

    1981-05-01

    The KUGEL computer code is designed to perform thermal/hydraulic analysis and coated-fuel particle performance calculations for axisymmetric pebble bed reactor (PBR) cores. This computer code was developed as part of a Department of Energy (DOE)-funded study designed to verify the published core performance data on PBRs. The KUGEL code is designed to interface directly with the 2DB code, a two-dimensional neutron diffusion code, to obtain distributions of thermal power, fission rate, fuel burnup, and fast neutron fluence, which are needed for thermal/hydraulic and fuel performance calculations. The code is variably dimensioned so that problem size can be easily varied. An interpolation routine allows variable mesh size to be used between the 2DB output and the two-dimensional thermal/hydraulic calculations.

  10. A preliminary assessment of the feasibility of deriving liquid and gaseous fuels from grown and waste organics

    NASA Technical Reports Server (NTRS)

    Graham, R. W.; Reynolds, T. W.; Hsu, Y.-Y.

    1976-01-01

    An estimate is obtained of the yearly supply of organic material for conversion to fuels, the energy potential is evaluated, and the fermentation and pyrolysis conversion processes are discussed. An investigation is conducted of the estimated cost of fuel from organics and the conclusions of an overall evaluation are presented. It is found that climate, land availability and economics of agricultural production and marketing, food demand, fertilizer shortage, and water availability combine to cast doubts on the feasibility of producing grown organic matter for fuel, in competition with food, feed, or fiber. Less controversial is the utilization of agricultural, industrial, and domestic waste as a conversion feedstock. The evaluation of a demonstration size system is recommended.

  11. Comprehensive modeling of a liquid rocket combustion chamber

    NASA Technical Reports Server (NTRS)

    Liang, P.-Y.; Fisher, S.; Chang, Y. M.

    1985-01-01

    An analytical model for the simulation of detailed three-phase combustion flows inside a liquid rocket combustion chamber is presented. The three phases involved are: a multispecies gaseous phase, an incompressible liquid phase, and a particulate droplet phase. The gas and liquid phases are continuum described in an Eulerian fashion. A two-phase solution capability for these continuum media is obtained through a marriage of the Implicit Continuous Eulerian (ICE) technique and the fractional Volume of Fluid (VOF) free surface description method. On the other hand, the particulate phase is given a discrete treatment and described in a Lagrangian fashion. All three phases are hence treated rigorously. Semi-empirical physical models are used to describe all interphase coupling terms as well as the chemistry among gaseous components. Sample calculations using the model are given. The results show promising application to truly comprehensive modeling of complex liquid-fueled engine systems.

  12. Japan's research on gaseous flames

    NASA Technical Reports Server (NTRS)

    Niioka, Takashi

    1995-01-01

    Although research studies on gaseous flames in microgravity in Japan have not been one-sided, they have been limited, for the most part, to comparatively fundamental studies. At present it is only possible to achieve a microgravity field by the use of drop towers, as far as gaseous flames are concerned. Compared with experiments on droplets, including droplet arrays, which have been vigorously performed in Japan, studies on gaseous flames have just begun. Experiments on ignition of gaseous fuel, flammability limits, flame stability, effect of magnetic field on flames, and carbon formation from gaseous flames are currently being carried out in microgravity. Seven subjects related to these topics are introduced and discussed herein.

  13. Robotic Manufacturing of 18-ft (5.5m) Diameter Cryogenic Fuel Tank Dome Assemblies for the NASA Ares I Rocket

    NASA Technical Reports Server (NTRS)

    Jones, Ronald E.; Carter, Robert W.

    2012-01-01

    The Ares I rocket was the first launch vehicle scheduled for manufacture under the National Aeronautic and Space Administration's Constellation program. A series of full-scale Ares I development articles were constructed on the Robotic Weld Tool at the NASA George C. Marshall Space Flight Center in Huntsville, Alabama. The Robotic Weld Tool is a 100 ton, 7- axis, robotic manufacturing system capable of machining and friction stir welding large-scale space hardware. This paper will focus on the friction stir welding of 18-ft (5.5m) diameter cryogenic fuel tank components; specifically, the liquid hydrogen forward dome and two common bulkhead manufacturing development articles.

  14. A Study of Pollutant Formation from the Lean Premixed Combustion of Gaseous Fuel Alternatives to Natural Gas

    NASA Astrophysics Data System (ADS)

    Fackler, Keith Boyd, Jr.

    The goal of this research is to identify how nitrogen oxide (NO x) emissions and flame stability (blowout) are impacted by the use of fuels that are alternatives to typical pipeline natural gas. The research focuses on lean, premixed combustors that are typically used in state-of-the-art natural gas fueled systems. An idealized laboratory lean premixed combustor, specifically the jet-stirred reactor, is used for experimental data. A series of models, including those featuring detailed fluid dynamics and those focusing on detailed chemistry, are used to interpret the data and understand the underlying chemical kinetic reasons for differences in emissions between the various fuel blends. An ultimate goal is to use these data and interpretive tools to develop a way to predict the emission and stability impacts of changing fuels within practical combustors. All experimental results are obtained from a high intensity, single-jet stirred reactor (JSR). Five fuel categories are studied: (1) pure H 2, (2) process and refinery gas, including combinations of H2, CH4, C2H6, and C3H8, (3) oxygen blown gasified coal/petcoke composed of H2, CO, and CO2, (4) landfill and digester gas composed of CH4, CO2, and N2, and (5) liquified natural gas (LNG)/shale/associated gases composed of CH4, C2H6, and C3 H8. NOx measurements are taken at a nominal combustion temperature of 1800 K, atmospheric pressure, and a reactor residence time of 3 ms. This is done to focus the results on differences caused by fuel chemistry by comparing all fuels at a common temperature, pressure, and residence time. This is one of the few studies in the literature that attempts to remove these effects when studying fuels varying in composition. Additionally, the effects of changing temperature and residence time are investigated for selected fuels. At the nominal temperature and residence time, the experimental and modeling results show the following trends for NOx emissions as a function of fuel type: 1.) NOx

  15. Early Rockets

    NASA Image and Video Library

    2004-04-15

    Sir William Congreve developed a rocket with a range of about 9,000 feet. The incendiary rocket used black powder, an iron case, and a 16-foot guide stick. In 1806, British used Congreve rockets to attack Napoleon's headquarters in France. In 1807, Congreve directed a rocket attack against Copenhagen.

  16. Early Rockets

    NASA Image and Video Library

    2004-04-15

    During the early introduction of rockets to Europe, they were used only as weapons. Enemy troops in India repulsed the British with rockets. Later, in Britain, Sir William Congreve developed a rocket that could fire to about 9,000 feet. The British fired Congreve rockets against the United States in the War of 1812.

  17. Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

    NASA Technical Reports Server (NTRS)

    Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

    2002-01-01

    This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

  18. Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

    NASA Technical Reports Server (NTRS)

    Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

    2002-01-01

    This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

  19. Early Rockets

    NASA Image and Video Library

    1940-01-01

    The Hermes A-1 rocket was designed by the U. S. Army after capturing the V-2 rocket from the German army at the conclusion of the Second World War. The Hermes A-1 is a modified V-2 rocket; it utilized the German aerodynamic configuration; however, internally it was a completely new design. This rocket was the first designed by the German Rocket Team at Redstone Arsenal in Huntsville, AL.

  20. Early Rockets

    NASA Image and Video Library

    2004-04-15

    Goddard rocket with four rocket motors. This rocket attained an altitude of 200 feet in a flight, November 1936, at Roswell, New Mexico. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

  1. Ignition Delay Experiments with Small-scale Rocket Engine at Simulated Altitude Conditions Using Various Fuels with Nitric Acid Oxidants / Dezso J. Ladanyi

    NASA Technical Reports Server (NTRS)

    Ladanyi, Dezso J

    1952-01-01

    Ignition delay determinations of several fuels with nitric oxidants were made at simulated altitude conditions utilizing a small-scale rocket engine of approximately 50 pounds thrust. Included in the fuels were aniline, hydrazine hydrate, furfuryl alcohol, furfuryl mercaptan, turpentine, and mixtures of triethylamine with mixed xylidines and diallyaniline. Red fuming, white fuming, and anhydrous nitric acids were used with and without additives. A diallylaniline - triethylamine mixture and a red fuming nitric acid analyzing 3.5 percent water and 16 percent NO2 by weight was found to have a wide temperature-pressure ignition range, yielding average delays from 13 milliseconds at 110 degrees F to 55 milliseconds at -95 degrees F regardless of the initial ambient pressure that ranged from sea-level pressure altitude of 94,000 feet.

  2. Focused RBCC Experiments: Two-Rocket Configuration Experiments and Hydrocarbon/Oxygen Rocket Ejector Experiments

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    This addendum report documents the results of two additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Penn State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3 d generation Reusable Launch Vehicles (RLV). The tasks reported here build on an earlier NASA MSFC funded research program on rocket ejector investigations. The first task investigated the improvements of a gaseous hydrogen/oxygen twin thruster RBCC rocket ejector system over a single rocket system. The second task investigated the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. Detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen, nitrogen and water vapor) for the gaseous hydrogen/oxygen rocket ejector experiments.

  3. A QSAR/QSTR Study on the Environmental Health Impact by the Rocket Fuel 1,1-Dimethyl Hydrazine and its Transformation Products

    PubMed Central

    Carlsen, Lars; Kenessov, Bulat N.; Batyrbekova, Svetlana Ye.

    2008-01-01

    QSAR/QSTR modelling constitutes an attractive approach to preliminary assessment of the impact on environmental health by a primary pollutant and the suite of transformation products that may be persistent in and toxic to the environment. The present paper studies the impact on environmental health by residuals of the rocket fuel 1,1-dimethyl hydrazine (heptyl) and its transformation products. The transformation products, comprising a variety of nitrogen containing compounds are suggested all to possess a significant migration potential. In all cases the compounds were found being rapidly biodegradable. However, unexpected low microbial activity may cause significant changes. None of the studied compounds appear to be bioaccumulating. Apart from substances with an intact hydrazine structure or hydrazone structure the transformation products in general display rather low environmental toxicities. Thus, it is concluded that apparently further attention should be given to tri- and tetramethyl hydrazine and 1-formyl 2,2-dimethyl hydrazine as well as to the hydrazones of formaldehyde and acetaldehyde as these five compounds may contribute to the overall environmental toxicity of residual rocket fuel and its transformation products. PMID:21572843

  4. A QSAR/QSTR Study on the Environmental Health Impact by the Rocket Fuel 1,1-Dimethyl Hydrazine and its Transformation Products.

    PubMed

    Carlsen, Lars; Kenessov, Bulat N; Batyrbekova, Svetlana Ye

    2008-07-18

    QSAR/QSTR modelling constitutes an attractive approach to preliminary assessment of the impact on environmental health by a primary pollutant and the suite of transformation products that may be persistent in and toxic to the environment. The present paper studies the impact on environmental health by residuals of the rocket fuel 1,1-dimethyl hydrazine (heptyl) and its transformation products. The transformation products, comprising a variety of nitrogen containing compounds are suggested all to possess a significant migration potential. In all cases the compounds were found being rapidly biodegradable. However, unexpected low microbial activity may cause significant changes. None of the studied compounds appear to be bioaccumulating.Apart from substances with an intact hydrazine structure or hydrazone structure the transformation products in general display rather low environmental toxicities. Thus, it is concluded that apparently further attention should be given to tri- and tetramethyl hydrazine and 1-formyl 2,2-dimethyl hydrazine as well as to the hydrazones of formaldehyde and acetaldehyde as these five compounds may contribute to the overall environmental toxicity of residual rocket fuel and its transformation products.

  5. Early Rockets

    NASA Image and Video Library

    2004-04-15

    The British fired Congreve rockets against the United States in the War of 1812. As a result Francis Scott Key coined the phrase the "rocket's red glare." Congreve had used a 16-foot guide stick to help stabilize his rocket. William Hale, another British inventor, invented the stickless rocket in 1846. The U.S. Army used the Hale rocket more than 100 years ago in the war with Mexico. Rockets were also used to a limited extent by both sides in the American Civil War.

  6. Congreve Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    The British fired Congreve rockets against the United States in the War of 1812. As a result Francis Scott Key coined the phrase the 'rocket's red glare.' Congreve had used a 16-foot guide stick to help stabilize his rocket. William Hale, another British inventor, invented the stickless rocket in 1846. The U.S. Army used the Hale rocket more than 100 years ago in the war with Mexico. Rockets were also used to a limited extent by both sides in the American Civil War.

  7. Congreve Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    The British fired Congreve rockets against the United States in the War of 1812. As a result Francis Scott Key coined the phrase the 'rocket's red glare.' Congreve had used a 16-foot guide stick to help stabilize his rocket. William Hale, another British inventor, invented the stickless rocket in 1846. The U.S. Army used the Hale rocket more than 100 years ago in the war with Mexico. Rockets were also used to a limited extent by both sides in the American Civil War.

  8. Experimental Study of Propulsion Performance by Single-Pulse Rotating Detonation with Gaseous Fuels-Oxygen Mixtures

    NASA Astrophysics Data System (ADS)

    Toshimitsu, Kazuhiko; Hara, Kosei; Mikajiri, Shuuto; Takiguchi, Naoki

    2016-12-01

    A rotating detonation engine (RDE) is one of candidates of aerospace engines for supersonic cruse, which is better for propulsion system than a pulse detonation engine (PDE) from the view of continuous thrust and simple structure. The propulsion performance of a proto-type RDE and a PDE by single pulse explosion with methane-oxygen is investigated. Furthermore, the performance of the RDE with acetylene-oxygen gas mixtures is investigated. Its impulse is estimated through ballistic pendulum method with maximum displacement and damping ratio. The comparison of specific impulses of the mixture gases at atmospheric pressure is shown. The specific impulses of the RDE and the PDE are almost same with methane-oxygen gas. Furthermore, the fuel-base specific impulse of the RDE with acetylene-oxygen gas is about over twice as large as one of methane-oxygen, and its maximum specific impulse is 1100 seconds.

  9. Early Rockets

    NASA Image and Video Library

    2004-04-15

    World War I enlisted rockets once again for military purposes. French pilots rigged rockets to the wing struts of their airplanes and aimed them at enemy observation balloons filled with highly inflammable hydrogen.

  10. Early Rockets

    NASA Image and Video Library

    2004-04-15

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  11. Rocket Flight.

    ERIC Educational Resources Information Center

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

    Describes an activity for designing, building, and launching rockets that provides students with an intrinsically motivating and real-life application of what could have been classroom-only concepts. Includes rocket design guidelines and a sample grading rubric. (KHR)

  12. Early Rockets

    NASA Image and Video Library

    2004-04-15

    In addition to Dr. Robert Goddard's pioneering work, American experimentation in rocketry prior to World War II grew, primarily in technical societies. This is an early rocket motor designed and developed by the American Rocket Society in 1932.

  13. Torpedo Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  14. Rocket Flight.

    ERIC Educational Resources Information Center

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

    Describes an activity for designing, building, and launching rockets that provides students with an intrinsically motivating and real-life application of what could have been classroom-only concepts. Includes rocket design guidelines and a sample grading rubric. (KHR)

  15. Torpedo Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  16. Evaluation of exhaust emissions from a Bi-fueled vehicle operating on liquid and gaseous fuels. Topical report, June-July 1995

    SciTech Connect

    Whitney, K.A.

    1995-12-01

    Exhaust emissions were characterized from a bi-fueled vehicle operating on compressed natural gas and two gasolines over a heavy acceleration/high speed driving cycle and during cold temperature operation. The test fuels included compressed natural gas (CNG) meeting California Air Resources Board emissions certification specifications, industry average gasoline formulated to the specifications of Reference Fuel A (RF-A) used in the CRC/Auto Oil program, and Federal reformulated gasoline (RFG) purchased at a commercial service station in metropolitan Houston. Exhaust emissions were evaluated over the light-duty chassis dynamometer portion of the Federal Test Procedure at 75 deg F and at 20 deg F, and the REP05 - a hot, stabilized, high speed, high acceleration driving cycle developed by the EPA to be representative of non-FTP, in-use driving. In addition, CNG emissions were evaluated over the US06 driving cycle. Average regulated exhaust emissions (total hydrocarbons, methane, carbon monoxide, and oxides of nitrogen) were evaluated in a manner consistent with the Code of Federal Regulations.

  17. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

  18. Early Rockets

    NASA Image and Video Library

    1940-01-01

    The cutaway drawing of the A-4 (Aggregate-4) rocket. Later renamed the V-2 (Vengeance Weapon-2), The rocket was developed by Dr. Wernher von Braun and the German rocket team at Peenemuende, Germany on the Baltic Sea. At the end of World War II, the team of German engineers and scientists came to the United States and continued rocket research for the Army at Fort Bliss, Texas, and Redstone Arsenal in Huntsville, Alabama.

  19. Small rocket flowfield diagnostic chambers

    NASA Technical Reports Server (NTRS)

    Morren, Sybil; Reed, Brian

    1993-01-01

    Instrumented and optically-accessible rocket chambers are being developed to be used for diagnostics of small rocket (less than 440 N thrust level) flowfields. These chambers are being tested to gather local fluid dynamic and thermodynamic flowfield data over a range of test conditions. This flowfield database is being used to better understand mixing and heat transfer phenomena in small rockets, influence the numerical modeling of small rocket flowfields, and characterize small rocket components. The diagnostic chamber designs include: a chamber design for gathering wall temperature profiles to be used as boundary conditions in a finite element heat flux model; a chamber design for gathering inner wall temperature and static pressure profiles; and optically-accessible chamber designs, to be used with a suite of laser-based diagnostics for gathering local species concentration, temperature, density, and velocity profiles. These chambers were run with gaseous hydrogen/gaseous oxygen (GH2/GO2) propellants, while subsequent versions will be run on liquid oxygen/hydrocarbon (LOX/HC) propellants. The purpose, design, and initial test results of these small rocket flowfield diagnostic chambers are summarized.

  20. Simultaneous high-speed schlieren and OH chemiluminescence imaging in a hybrid rocket combustor at elevated pressures

    NASA Astrophysics Data System (ADS)

    Miller, Victor; Jens, Elizabeth T.; Mechentel, Flora S.; Cantwell, Brian J.; Stanford Propulsion; Space Exploration Group Team

    2014-11-01

    In this work, we present observations of the overall features and dynamics of flow and combustion in a slab-type hybrid rocket combustor. Tests were conducted in the recently upgraded Stanford Combustion Visualization Facility, a hybrid rocket combustor test platform capable of generating constant mass-flux flows of oxygen. High-speed (3 kHz) schlieren and OH chemiluminescence imaging were used to visualize the flow. We present imaging results for the combustion of two different fuel grains, a classic, low regression rate polymethyl methacrylate (PMMA), and a high regression rate paraffin, and all tests were conducted in gaseous oxygen. Each fuel grain was tested at multiple free-stream pressures at constant oxidizer mass flux (40 kg/m2s). The resulting image sequences suggest that aspects of the dynamics and scaling of the system depend strongly on both pressure and type of fuel.

  1. 2D and 3D Modeling Efforts in Fuel Film Cooling of Liquid Rocket Engines (Conference Paper with Briefing Charts)

    DTIC Science & Technology

    2017-01-12

    barrier coating, insulating the wall as the fuel film begins to dissipate. Another important physical factor could be radiative heat transfer, which...important to capture these physical effects to better characterize the effectiveness of a given fuel film cooling (FFC) geometry. This paper covers...the parametric studies also serve to provide a better understanding of the physics driving the unsteadiness in a fuel-film cooled configuration. This

  2. Mixing, chemical reaction and flow field development in ducted rockets

    SciTech Connect

    Vanka, S.P.; Craig, R.R.; Stull, F.D.

    1984-09-01

    Calculations have been made of the three-dimensional mixing, chemical reaction, and flow field development in a typical ducted rocket configuration. The governing partial differential equations are numerically solved by an iterative finite-difference solution procedure. The physical models include the k approx. epsilon turbulence model, one-step reaction, and mixing controlled chemical reaction rate. Radiation is neglected. The mean flow structure, fuel dispersal patterns, and temperature field are presented in detail for a base configuration with 0.058 m (2 in.) dome height, 45/sup 0/ side arm inclination, and with gaseous ethylene injected from the dome plate at an eccentric location. In addition, the influences of the geometrical parameters such as dome height, inclination of the side arms, and location of the fuel injector are studied.

  3. Early Rockets

    NASA Image and Video Library

    2004-04-15

    During the 19th century, rocket enthusiasts and inventors began to appear in almost every country. Some people thought these early rocket pioneers were geniuses, and others thought they were crazy. Claude Ruggieri, an Italian living in Paris, apparently rocketed small animals into space as early as 1806. The payloads were recovered by parachute. As depicted here by artist Larry Toschik, French authorities were not always impressed with rocket research. They halted Ruggieri's plans to launch a small boy using a rocket cluster. (Reproduced from a drawing by Larry Toschik and presented here courtesy of the artist and Motorola Inc.)

  4. Driving of combustor oscillations by gaseous propellant injectors

    NASA Technical Reports Server (NTRS)

    Janardan, B. A.; Daniel, B. R.; Zinn, B. T.

    1979-01-01

    Measurements of reactive admittances that describe the capabilities of the combustion processes associated with coaxial gaseous fuel injectors to amplify combustor oscillations are presented. These admittances are needed in (1) stability analyses of rocket motors and gas turbine combustors; (2) for the evaluation of 'driving' capabilities of injectors; and (3) for checking the applicability of a theoretical model. The modified standing-wave technique was used to determine the admittances of the combustion processes in coaxial injectors utilizing air-acetylene mixtures. The measurements indicated that (1) coaxial injectors can sustain combustion instabilities over wide frequency ranges; (2) the maximum driving capability of an injector decreases with an increased equivalence ratio; (3) the frequency at which maximum driving is observed decreases with the increased equivalence ratio; and (4) the characteristic combustion time of an injector design decreases with increased frequency.

  5. Durability Demonstration of Small High Performance Rocket Thrusters for RLV Applications

    NASA Technical Reports Server (NTRS)

    Neill, Todd; Gladstone, Jeffrey

    1999-01-01

    The development of low cost, highly reusable rocket thrusters is essential for fulfilling NASA's commitment of providing economical space access. Results obtained from a recent sub-scale, high performance rocket thruster development and test program applied to a Rocket Based Combined Cycle engine have demonstrated rocket versatility and durability in over 678 tests at chamber pressures between 190 and 1930 psia using gaseous oxygen and gaseous hydrogen.

  6. Experimental/Analytical Characterization of the RBCC Rocket-Ejector Mode

    NASA Technical Reports Server (NTRS)

    Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.

    2000-01-01

    The experimental/analytical research work described here addresses the rocket-ejector mode (Mach 0-2 operational range) of the RBCC engine. The experimental phase of the program includes studying the mixing and combustion characteristics of the rocket-ejector system utilizing state-of-the-art diagnostic techniques. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was utilized as the experimental platform. The goals of the experimental phase of the research being conducted at Penn State are to: (a) systematically increase the range of rocket-ejector understanding over a wide range of flow/geometry parameters and (b) provide a comprehensive data base for evaluating and anchoring CFD codes. Concurrent with the experimental activities, a CFD code benchmarking effort at Marshall Space Flight Center is also being used to further investigate the RBCC rocket-ejector mode. Experiments involving the single rocket based optically-accessible rocket-ejector system have been conducted for Diffusion and Afterburning (DAB) as well as Simultaneous Mixing and Combustion configurations. For the DAB configuration, air is introduced (direct-connect) or ejected (sea-level static) into a constant area mixer section with a centrally located gaseous oxygen (GO2)/gaseous hydrogen (GH2) rocket combustor. The downstream flowpath for this configuration includes a diffuser, an afterburner and a final converging nozzle. For the SMC configuration, the rocket is centrally located in a slightly divergent duct. For all tested configurations, global measurements of the axial pressure and heat transfer profiles as well as the overall engine thrust were made. Detailed measurements include major species concentration (H2 O2 N2 and H2O) profiles at various mixer locations made using Raman spectroscopy. Complementary CFD calculations of the flowfield at the experimental conditions also provide additional information on the physics of the problem. These calculations

  7. 14 CFR 34.71 - Compliance with gaseous emission standards.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... TRANSPORTATION AIRCRAFT FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE ENGINE POWERED AIRPLANES Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft and Aircraft Gas Turbine Engines) §...

  8. 14 CFR 34.71 - Compliance with gaseous emission standards.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... TRANSPORTATION AIRCRAFT FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE ENGINE POWERED AIRPLANES Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft and Aircraft Gas Turbine Engines) §...

  9. Focused Rocket-Ejector RBCC Experiments

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    This document reports the results of additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Perm State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3rd generation Reusable Launch Vehicles (RLV). The two tasks conducted under this program build on earlier NASA MSFC funded research program on rocket ejector investigations. The first task continued a systematic investigation of the improvements provided by a gaseous hydrogen (GHz)/oxygen (GO2) twin thruster RBCC rocket ejector system over a single rocket system. In a similar vein, the second task continued investigations into the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. For the GH2/GO2 propellant rocket ejector experiments, high frequency measurements of the pressure field within the system were also made to understand the unsteady behavior of the flowfield.

  10. Design and evaluation of an oxidant-fuel-ratio-zoned rocket injector for high performance and ablative engine compatibility

    NASA Technical Reports Server (NTRS)

    Winter, J. M.; Pavli, A. J.; Shinn, A. M., Jr.

    1972-01-01

    A method for temperature control of the combustion gases in the peripheral zone of a rocket combustor which would reduce ablative throat erosion, prevent melting of zirconia throat inserts, and maintain high combustion performance is discussed. Included are techniques for analyzing and predicting zoned injector performance, as well as the philosophy and method for accomplishing an optimum compromise between high performance and reduced effective gas temperature. The experimental work was done with a 1000-lbf rocket engine which used as propellants N2O4 and a blend of 50-percent N2H4 and 50-percent UDMH at 100-psia chamber pressure and an overall O/F of 2.0. The method selected to provide temperature control was to use 30 percent of the propellant to form a peripheral zone of combustion gases at an O/F of 1.31 and 2700 K. The remaining 70 percent of the propellant in the core was at an O/F of 2.45 to keep the overall O/F at 2.0.

  11. Early Rockets

    NASA Image and Video Library

    1958-01-31

    Juno I, a slightly modified Jupiter-C launch vehicle, shortly before the January 31, 1958 launch of America's first satellite, Explorer I. The Jupiter-C, developed by Dr. Wernher von Braun and the rocket team at Redstone Arsenal in Huntsville, Alabama, consisted of a modified version of the Redstone rocket's first stage and two upper stages of clustered Baby Sergeant rockets developed by the Jet Propulsion Laboratory.

  12. Early Rockets

    NASA Image and Video Library

    1940-01-01

    This German cutaway drawing of the Aggregate-4 (A-4) illustrates the dimensions and internal workings of the rocket. Later renamed the V-2, the rocket was developed by Dr. Wernher von Braun and the German Rocket Team at Peenemuende on the Baltic Sea. At the end of World War II, the team of German engineers and scientists came to the United States to work for the Army at Fort Bliss, Texas, and Redstone Arsenal in Huntsville, Alabama.

  13. Early Rockets

    NASA Image and Video Library

    1940-01-01

    This drawing illustrates the vital dimensions of the A-4 (Aggregate-4). Later renamed the V-2 (Vengeance Weapon-2), the rocket was developed by Dr. Wernher von Braun and the German rocket team at Peenemuende, Germany on the Baltic Sea. At the end of World War II, the team of German engineers and scientists came to the United States and continued rocket research for the Army at Fort Bliss, Texas, and Redstone Arsenal in Huntsville, Alabama.

  14. Experimental Analysis of a Rocket Based Combined Cycle (RBCC) Engine in a Direct-Connect Test Facility

    NASA Technical Reports Server (NTRS)

    Nelson, K.; Hawk, Clark W.

    1997-01-01

    The object of this study is to investigate the operation of a RBCC at ramjet and scramjet flight conditions using a direct-connect test facility. The apparatus being tested is a single strut-rocket within a dual-mode ram/scramjet combustor. The gaseous hydrogen/oxygen, linear strut-rocket was supplied by Aerojet Propulsion Company. The hardware is being tested in the Direct Connect Supersonic Combustion Test Facility at NASA Langley Research Center. The test facilities hydrogen/oxygen vitiated heater is capable of flight total enthalpies to Mach 8. A Mach 2.5 facility nozzle mates the heater to the combustor duct. The rocket ejector will ordinarily operate in a fuel-rich mode. Additional fuel injection is provided by a pair of parallel injectors located at the base of the strut body. Instrumentation on the test apparatus includes a unique, direct thrust measurement system. Performance predictions for the anticipated test conditions have been made using a one-dimensional, thermodynamic analysis code. Results from the code show the dependence of overall thrust and specific impulse on rocket chamber pressure, rocket fuel equivalence ratio, and overall fuel equivalence ratio. Once the experimental test series begins, the inferred combustion efficiency as a function of axial location and the thermal choke region (where applicable) can also be determined using this code. Upon completion of the experimental test series, measurements will be used to calculate thrust, specific impulse, etc. Measured and calculated values will be compared to those found analytically. If appropriate, the code will be tailored to better predict hardware operation. Conclusions will be drawn as to the fuel-rich rocket's overall effect on ramjet and scramjet performance. Also, comparisons will be made between the integrated thrust calculated from the static pressure taps located along the duct and the thrust measured by the direct thrust measurement system.

  15. Experimental Analysis of a Rocket Based Combined Cycle (RBCC) Engine in a Direct-Connect Test Facility

    NASA Technical Reports Server (NTRS)

    Nelson, K.; Hawk, Clark W.

    1997-01-01

    The object of this study is to investigate the operation of a RBCC at ramjet and scramjet flight conditions using a direct-connect test facility. The apparatus being tested is a single strut-rocket within a dual-mode ram/scramjet combustor. The gaseous hydrogen/oxygen, linear strut-rocket was supplied by Aerojet Propulsion Company. The hardware is being tested in the Direct Connect Supersonic Combustion Test Facility at NASA Langley Research Center. The test facilities hydrogen/oxygen vitiated heater is capable of flight total enthalpies to Mach 8. A Mach 2.5 facility nozzle mates the heater to the combustor duct. The rocket ejector will ordinarily operate in a fuel-rich mode. Additional fuel injection is provided by a pair of parallel injectors located at the base of the strut body. Instrumentation on the test apparatus includes a unique, direct thrust measurement system. Performance predictions for the anticipated test conditions have been made using a one-dimensional, thermodynamic analysis code. Results from the code show the dependence of overall thrust and specific impulse on rocket chamber pressure, rocket fuel equivalence ratio, and overall fuel equivalence ratio. Once the experimental test series begins, the inferred combustion efficiency as a function of axial location and the thermal choke region (where applicable) can also be determined using this code. Upon completion of the experimental test series, measurements will be used to calculate thrust, specific impulse, etc. Measured and calculated values will be compared to those found analytically. If appropriate, the code will be tailored to better predict hardware operation. Conclusions will be drawn as to the fuel-rich rocket's overall effect on ramjet and scramjet performance. Also, comparisons will be made between the integrated thrust calculated from the static pressure taps located along the duct and the thrust measured by the direct thrust measurement system.

  16. Early Rockets

    NASA Image and Video Library

    1926-03-16

    Dr. Goddard's 1926 rocket configuration. Dr. Goddard's liquid oxygen-gasoline rocket was fired on March 16, 1926, at Auburn, Massachusetts. It flew for only 2.5 seconds, climbed 41 feet, and landed 184 feet away in a cabbage patch. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets, which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

  17. Early Rockets

    NASA Image and Video Library

    2004-04-15

    By 1870, American and British inventors had found other ways to use rockets. For example, the Congreve rocket was capable of carrying a line over 1,000 feet to a stranded ship. In 1914, an estimated 1,000 lives were saved by this technique.

  18. Early Rockets

    NASA Image and Video Library

    2004-04-15

    One of the earliest recorded instances of the use of rockets was as military weapons against the Mongols by the Chinese at the siege of Kai Fung Foo in 1232 A.D. An arrow with a tube of gunpowder produced an arrow of flying fire. The Mongol attackers fled in terror, even though the rockets were inaccurate and relatively harmless.

  19. Early Rockets

    NASA Image and Video Library

    2004-04-15

    In the 19th Century, experiments in America, Europe, and elsewhere attempted to build postal rockets to deliver mail from one location to another. The idea was more novel than successful. Many stamps used in these early postal rockets have become collector's items.

  20. Early Rockets

    NASA Image and Video Library

    1947-01-01

    A V-2 rocket is hoisted into a static test facility at White Sands, New Mexico. The German engineers and scientists who developed the V-2 came to the United States at the end of World War II and continued rocket testing under the direction of the U. S. Army, launching more than sixty V-2s.

  1. Early Rockets

    NASA Image and Video Library

    1946-01-01

    A V-2 rocket takes flight at White Sands, New Mexico, in 1946. The German engineers and scientists who developed the V-2 came to the United States at the end of World War II and continued rocket testing under the direction of the U. S. Army, launching more than sixty V-2s.

  2. Gaseous Detectors

    NASA Astrophysics Data System (ADS)

    Titov, Maxim

    Since long time, the compelling scientific goals of future high-energy physics experiments were a driving factor in the development of advanced detector technologies. A true innovation in detector instrumentation concepts came in 1968, with the development of a fully parallel readout for a large array of sensing elements - the Multi-Wire Proportional Chamber (MWPC), which earned Georges Charpak a Nobel prize in physics in 1992. Since that time radiation detection and imaging with fast gaseous detectors, capable of economically covering large detection volumes with low mass budget, have been playing an important role in many fields of physics. Advances in photolithography and microprocessing techniques in the chip industry during the past decade triggered a major transition in the field of gas detectors from wire structures to Micro-Pattern Gas Detector (MPGD) concepts, revolutionizing cell-size limitations for many gas detector applications. The high radiation resistance and excellent spatial and time resolution make them an invaluable tool to confront future detector challenges at the next generation of colliders. The design of the new micro-pattern devices appears suitable for industrial production. Novel structures where MPGDs are directly coupled to the CMOS pixel readout represent an exciting field allowing timing and charge measurements as well as precise spatial information in 3D. Originally developed for the high-energy physics, MPGD applications have expanded to nuclear physics, photon detection, astroparticle and neutrino physics, neutron detection, and medical imaging.

  3. Rocket-in-a-Duct Performance Analysis

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Reed, Brian D.

    1999-01-01

    An axisymmetric, 110 N class, rocket configured with a free expansion between the rocket nozzle and a surrounding duct was tested in an altitude simulation facility. The propellants were gaseous hydrogen and gaseous oxygen and the hardware consisted of a heat sink type copper rocket firing through copper ducts of various diameters and lengths. A secondary flow of nitrogen was introduced at the blind end of the duct to mix with the primary rocket mass flow in the duct. This flow was in the range of 0 to 10% of the primary massflow and its effect on nozzle performance was measured. The random measurement errors on thrust and massflow were within +/-1%. One dimensional equilibrium calculations were used to establish the possible theoretical performance of these rocket-in-a-duct nozzles. Although the scale of these tests was small, they simulated the relevant flow expansion physics at a modest experimental cost. Test results indicated that lower performance was obtained at higher free expansion area ratios and longer ducts, while, higher performance was obtained with the addition of secondary flow. There was a discernable peak in specific impulse efficiency at 4% secondary flow. The small scale of these tests resulted in low performance efficiencies, but prior numerical modeling of larger rocket-in-a-duct engines predicted performance that was comparable to that of optimized rocket nozzles. This remains to be proven in large-scale, rocket-in-a-duct tests.

  4. Nuclear Rocket Technology Conference

    NASA Technical Reports Server (NTRS)

    1966-01-01

    The Lewis Research Center has a strong interest in nuclear rocket propulsion and provides active support of the graphite reactor program in such nonnuclear areas as cryogenics, two-phase flow, propellant heating, fluid systems, heat transfer, nozzle cooling, nozzle design, pumps, turbines, and startup and control problems. A parallel effort has also been expended to evaluate the engineering feasibility of a nuclear rocket reactor using tungsten-matrix fuel elements and water as the moderator. Both of these efforts have resulted in significant contributions to nuclear rocket technology. Many successful static firings of nuclear rockets have been made with graphite-core reactors. Sufficient information has also been accumulated to permit a reasonable Judgment as to the feasibility of the tungsten water-moderated reactor concept. We therefore consider that this technoIogy conference on the nuclear rocket work that has been sponsored by the Lewis Research Center is timely. The conference has been prepared by NASA personnel, but the information presented includes substantial contributions from both NASA and AEC contractors. The conference excludes from consideration the many possible mission requirements for nuclear rockets. Also excluded is the direct comparison of nuclear rocket types with each other or with other modes of propulsion. The graphite reactor support work presented on the first day of the conference was partly inspired through a close cooperative effort between the Cleveland extension of the Space Nuclear Propulsion Office (headed by Robert W. Schroeder) and the Lewis Research Center. Much of this effort was supervised by Mr. John C. Sanders, chairman for the first day of the conference, and by Mr. Hugh M. Henneberry. The tungsten water-moderated reactor concept was initiated at Lewis by Mr. Frank E. Rom and his coworkers. The supervision of the recent engineering studies has been shared by Mr. Samuel J. Kaufman, chairman for the second day of the

  5. Exploring Sustainable Rocket Fuels: [Imidazolyl-Amine-BH2](+)-Cation-Based Ionic Liquids as Replacements for Toxic Hydrazine Derivatives.

    PubMed

    Huang, Shi; Qi, Xiujuan; Zhang, Wenquan; Liu, Tianlin; Zhang, Qinghua

    2015-12-01

    The application of hypergolic ionic liquids as propellant fuels is a newly emerging area in the fields of chemistry and propulsion science. Herein, a new class of [imidazolyl-amine-BH2](+)-cation-based ionic liquids, which included fuel-rich anions, such as dicyanamide (N(CN)2(-)) and cyanoborohydride (BH3CN(-)) anions, were synthesized and characterized. As expected, all of the ionic liquids exhibited spontaneous combustion upon contact with the oxidizer 100 % HNO3. The densities of these ionic liquids varied from 0.99-1.12 g cm(-3), and the heats of formation, predicted based on Gaussian 09 calculations, were between -707.7 and 241.8 kJ mol(-1). Among them, the salt of compound 5, that is, (1-allyl-1H-imidazole-3-yl)-(trimethylamine)-dihydroboronium dicyanamide, exhibited the lowest viscosity (168 MPa s), good thermal properties (Tg <-70 °C, Td >130 °C), and the shortest ignition-delay time (18 ms) with 100 % HNO3. These ionic fuels, as "green" replacements for toxic hydrazine-derivatives, may have potential applications as bipropellant formulations.

  6. Early Rockets

    NASA Image and Video Library

    1926-03-16

    Dr. Robert H. Goddard and liquid oxygen-gasoline rocket in the frame from which it was fired on March 16, 1926, at Auburn, Mass. It flew for only 2.5 seconds, climbed 41 feet, and landed 184 feet away in a cabbage patch. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets, which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

  7. Rockets Away!

    ERIC Educational Resources Information Center

    Kaahaaina, Nancy

    1997-01-01

    Describes a project that involved a rocket-design competition where students played the roles of McDonnell Douglas employees competing for NASA contracts. Provides a real world experience involving deadlines, design and performance specifications, and budgets. (JRH)

  8. Early Rockets

    NASA Image and Video Library

    1958-01-31

    Launch of Jupiter-C/Explorer 1 at Cape Canaveral, Florida on January 31, 1958. After the Russian Sputnik 1 was launched in October 1957, the launching of an American satellite assumed much greater importance. After the Vanguard rocket exploded on the pad in December 1957, the ability to orbit a satellite became a matter of national prestige. On January 31, 1958, slightly more than four weeks after the launch of Sputnik.The ABMA (Army Ballistic Missile Agency) in Redstone Arsenal, Huntsville, Alabama, in cooperation with the Jet Propulsion Laboratory, launched a Jupiter from Cape Canaveral, Florida. The rocket consisted of a modified version of the Redstone rocket's first stage and two upper stages of clustered Baby Sergeant rockets developed by the Jet Propulsion Laboratory and later designated as Juno boosters for space launches

  9. Rockets Away!

    ERIC Educational Resources Information Center

    Kaahaaina, Nancy

    1997-01-01

    Describes a project that involved a rocket-design competition where students played the roles of McDonnell Douglas employees competing for NASA contracts. Provides a real world experience involving deadlines, design and performance specifications, and budgets. (JRH)

  10. Recent work on gaseous detonations

    NASA Astrophysics Data System (ADS)

    Nettleton, M. A.

    The paper reviews recent progress in the field of gaseous detonations, with sections on shock diffraction and reflection, the transition to detonation, hybrid, spherically-imploding, and galloping and stuttering fronts, their structure, their transmission and quenching by additives, the critical energy for initiation and detonation of more unusual fuels. The final section points out areas where our understanding is still far from being complete and contains some suggestions of ways in which progress might be made.

  11. Early Rockets

    NASA Image and Video Library

    1953-08-30

    U.S. Army Redstone Rocket: The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone rocket was also known as "Old Reliable" because of its many diverse missions. The first Redstone Missile was launched from Cape Canaveral, Florida on August 30, 1953.

  12. Early Rockets

    NASA Image and Video Library

    1957-10-03

    America’s first scientific satellite, the Explorer I, carried the radiation detection experiment designed by Dr. James Van Allen and discovered the Van Allen Radiation Belt. It was launched aboard a modified redstone rocket known as the Jupiter C, developed by Dr. von Braun’s rocket team at Redstone Arsenal in Huntsville, Alabama. The satellite launched on January 31, 1958, just 3 months after the the von Braun team received the go-ahead.

  13. Early Rockets

    NASA Image and Video Library

    1958-01-31

    Explorer 1 atop a Jupiter-C in gantry. Jupiter-C carrying the first American satellite, Explorer 1, was successfully launched on January 31, 1958. The Jupiter-C launch vehicle consisted of a modified version of the Redstone rocket's first stage and two upper stages of clustered Baby Sergeant rockets developed by the Jet Propulsion Laboratory and later designated as Juno boosters for space launches

  14. Early Rockets

    NASA Image and Video Library

    1950-02-24

    Bumper Wac liftoff at the Long Range Proving Ground located at Cape Canaveral, Florida. At White Sands, New Mexico, the German rocket team experimented with a two-stage rocket called Bumper Wac, which intended to provide data for upper atmospheric research. On February 24, 1950, the Bumper, which employed a V-2 as the first stage with a Wac Corporal upper stage, obtained a peak altitude of more than 240 miles.

  15. Early Rockets

    NASA Image and Video Library

    2004-04-15

    Dr. Robert H. Goddard loading a 1918 version of the Bazooka of World War II. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets, which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

  16. Multiple dopant injection system for small rocket engines

    NASA Technical Reports Server (NTRS)

    Sakala, G. G.; Raines, N. G.

    1992-01-01

    The Diagnostics Test Facility (DTF) at NASA's Stennis Space Center (SSC) was designed and built to provide a standard rocket engine exhaust plume for use in the research and development of engine health monitoring instrumentation. A 1000 lb thrust class liquid oxygen (LOX)-gaseous hydrogen (GH2) fueled rocket engine is used as the subscale plume source to simulate the SSME during experimentation and instrument development. The ability of the DTF to provide efficient, and low cost test operations makes it uniquely suited for plume diagnostic experimentation. The most unique feature of the DTF is the Multiple Dopant Injection System (MDIS) that is used to seed the exhaust plume with the desired element or metal alloy. The dopant injection takes place at the fuel injector, yielding a very uniform and homogeneous distribution of the seeding material in the exhaust plume. The MDIS allows during a single test firing of the DTF, the seeding of the exhaust plume with up to three different dopants and also provides distilled water base lines between the dopants. A number of plume diagnostic-related experiments have already utilized the unique capabilities of the DTF.

  17. Multiple dopant injection system for small rocket engines

    NASA Astrophysics Data System (ADS)

    Sakala, G. G.; Raines, N. G.

    1992-07-01

    The Diagnostics Test Facility (DTF) at NASA's Stennis Space Center (SSC) was designed and built to provide a standard rocket engine exhaust plume for use in the research and development of engine health monitoring instrumentation. A 1000 lb thrust class liquid oxygen (LOX)-gaseous hydrogen (GH2) fueled rocket engine is used as the subscale plume source to simulate the SSME during experimentation and instrument development. The ability of the DTF to provide efficient, and low cost test operations makes it uniquely suited for plume diagnostic experimentation. The most unique feature of the DTF is the Multiple Dopant Injection System (MDIS) that is used to seed the exhaust plume with the desired element or metal alloy. The dopant injection takes place at the fuel injector, yielding a very uniform and homogeneous distribution of the seeding material in the exhaust plume. The MDIS allows during a single test firing of the DTF, the seeding of the exhaust plume with up to three different dopants and also provides distilled water base lines between the dopants. A number of plume diagnostic-related experiments have already utilized the unique capabilities of the DTF.

  18. Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

    NASA Technical Reports Server (NTRS)

    Meyer, Mike L.

    1993-01-01

    Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel.

  19. Advanced bioreactor systems for gaseous substrates: Conversion of synthesis gas to liquid fuels and removal of SO{sub X} and NO{sub X} from coal combustion gases

    SciTech Connect

    Selvaraj, P.T.; Kaufman, E.N.

    1996-06-01

    The purpose of this research program is the development and demonstration of a new generation of gaseous substrate based bioreactors for the production of liquid fuels from coal synthesis gas and the removal of NO{sub x} and SO{sub x} species from combustion flue gas. This R&D program is a joint effort between the staff of the Bioprocessing Research and Development Center (BRDC) of ORNL and the staff of Bioengineering Resources, Inc. (BRI) under a Cooperative Research and Development Agreement (CRADA). The Federal Coordinating Council for Science, Engineering, and Technology report entitled {open_quotes}Biotechnology for the 21st Century{close_quotes} and the recent Energy Policy Act of 1992 emphasizes research, development, and demonstration of the conversion of coal to gaseous and liquid fuels and the control of sulfur and nitrogen oxides in effluent streams. This R&D program presents an innovative approach to the use of bioprocessing concepts that will have utility in both of these identified areas.

  20. Advanced bioreactor concepts for gaseous substrates: Conversion of synthesis gas to liquid fuels and removal of SO{sub x} and NO{sub x} from coal combustion gases. CRADA final report

    SciTech Connect

    Kaufman, E.N.; Selvaraj, P.T.

    1997-10-01

    The purpose of the proposed research program was the development and demonstration of a new generation of gaseous substrate-based bioreactors for the production of liquid fuels from coal synthesis gas and the removal of NO{sub x} and SO{sub x} species from coal combustion flue gas. This study addressed the further investigation of optimal bacterial strains, growth media and kinetics for the biocatalytic conversion of coal synthesis gas to liquid fuel such as ethanol and the reduction of gaseous flue gas constituents. The primary emphasis was on the development of advanced bioreactor systems coupled with innovative biocatalytic systems that will provide increased productivity under controlled conditions. It was hoped that this would result in bioprocessing options that have both technical and economic feasibility, thus, ensuring early industrial use. Predictive mathematical models were formulated to accommodate hydrodynamics, mass transport, and conversion kinetics, and provide the data base for design and scale-up. The program was separated into four tasks: (1) Optimization of Biocatalytic Kinetics; (2) Development of Well-mixed and Columnar Reactors; (3) Development of Predictive Mathematical Models; and (4) Industrial Demonstration. Research activities addressing both synthesis gas conversion and flue gas removal were conducted in parallel by BRI and ORNL respectively.

  1. Air-Breathing Rocket Engines

    NASA Technical Reports Server (NTRS)

    1998-01-01

    This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  2. Correlation of rocket propulsion fuel properties with chemical composition using comprehensive two-dimensional gas chromatography with time-of-flight mass spectrometry followed by partial least squares regression analysis

    SciTech Connect

    Kehimkar, Benjamin; Hoggard, Jamin C.; Marney, Luke C.; Billingsley, Matthew; Fraga, Carlos G.; Bruno, Thomas J.; Synovec, Robert E.

    2014-01-31

    There is an increased need to more fully assess and control the composition of kerosene based rocket propulsion fuels, namely RP-1 and RP-2. In particular, it is crucial to be able to make better quantitative connections between the following three attributes: (a) fuel performance, (b) fuel properties (flash point, density, kinematic viscosity, net heat of combustion, hydrogen content, etc) and (c) the chemical composition of a given fuel (i.e., specific chemical compounds and compound classes present as a result of feedstock blending and processing). Indeed, recent efforts in predicting fuel performance through modeling put greater emphasis on detailed and accurate fuel properties and fuel compositional information. In this regard, advanced distillation curve (ADC) metrology provides improved data relative to classical boiling point and volatility curve techniques. Using ADC metrology, data obtained from RP-1 and RP-2 fuels provides compositional variation information that is directly relevant to predictive modeling of fuel performance. Often, in such studies, one-dimensional gas chromatography (GC) combined with mass spectrometry (MS) is typically employed to provide chemical composition information. Building on approaches using GC-MS, but to glean substantially more chemical composition information from these complex fuels, we have recently studied the use of comprehensive two dimensional gas chromatography combined with time-of-flight mass spectrometry (GC × GC - TOFMS) to provide chemical composition data that is significantly richer than that provided by GC-MS methods. In this report, by applying multivariate data analysis techniques, referred to as chemometrics, we are able to readily model (correlate) the chemical compositional information from RP-1 and RP-2 fuels provided using GC × GC - TOFMS, to the fuel property information such as that provided by the ADC method and other specification properties. We anticipate that this new chemical analysis

  3. Focused Experimental and Analytical Studies of the RBCC Rocket-Ejector

    NASA Technical Reports Server (NTRS)

    Lehman, M.; Pal, S.; Schwes, D.; Chen, J. D.; Santoro, R. J.

    1999-01-01

    The rocket-ejector mode of a Rocket Based Combined Cycle Engine (RBCC) was studied through a joint experimental/analytical approach. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was designed and fabricated for experimentation. The rocket-ejector system utilizes a single two-dimensional gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a systematic understanding of the rocket ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions Overall system performance was obtained through Global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen. nitrogen and water vapor). These experimental efforts were complemented by Computational Fluid Dynamic (CFD) flowfield analyses.

  4. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

    Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

  5. Correlation of rocket propulsion fuel properties with chemical composition using comprehensive two-dimensional gas chromatography with time-of-flight mass spectrometry followed by partial least squares regression analysis.

    PubMed

    Kehimkar, Benjamin; Hoggard, Jamin C; Marney, Luke C; Billingsley, Matthew C; Fraga, Carlos G; Bruno, Thomas J; Synovec, Robert E

    2014-01-31

    There is an increased need to more fully assess and control the composition of kerosene-based rocket propulsion fuels such as RP-1. In particular, it is critical to make better quantitative connections among the following three attributes: fuel performance (thermal stability, sooting propensity, engine specific impulse, etc.), fuel properties (such as flash point, density, kinematic viscosity, net heat of combustion, and hydrogen content), and the chemical composition of a given fuel, i.e., amounts of specific chemical compounds and compound classes present in a fuel as a result of feedstock blending and/or processing. Recent efforts in predicting fuel chemical and physical behavior through modeling put greater emphasis on attaining detailed and accurate fuel properties and fuel composition information. Often, one-dimensional gas chromatography (GC) combined with mass spectrometry (MS) is employed to provide chemical composition information. Building on approaches that used GC-MS, but to glean substantially more chemical information from these complex fuels, we recently studied the use of comprehensive two dimensional (2D) gas chromatography combined with time-of-flight mass spectrometry (GC×GC-TOFMS) using a "reversed column" format: RTX-wax column for the first dimension, and a RTX-1 column for the second dimension. In this report, by applying chemometric data analysis, specifically partial least-squares (PLS) regression analysis, we are able to readily model (and correlate) the chemical compositional information provided by use of GC×GC-TOFMS to RP-1 fuel property information such as density, kinematic viscosity, net heat of combustion, and so on. Furthermore, we readily identified compounds that contribute significantly to measured differences in fuel properties based on results from the PLS models. We anticipate this new chemical analysis strategy will have broad implications for the development of high fidelity composition-property models, leading to an

  6. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust. The test was the first test ever anywhere outside Russia of a Russian designed and built engine.

  7. Aircraft Rockets,

    DTIC Science & Technology

    1981-10-16

    largest power stations, which lead industrial arsas, etc. The prominent representatives cf American and English VVS confirm that strategic aviatic best...rocket engine; IRD - turbcjet angine; * - no data . REFERENCES : B H. 1VeA0oCbe, r. B. CHHt up es. BeeAere a paeTHylo TexHHKy. OdoporH, 1956. A. C. A 0 K K...UNEDITED MACHINE TRANSLATION FTD-ID(RS)T-1954-80 J 16 Octobe’ 1981 MICROFICHE NR: FTD-81-C-000931 ( . IRCRAFT ROCKETS) By/>B. T./Surikov English pages: 138

  8. Early Rockets

    NASA Image and Video Library

    2004-04-15

    Engine for the Jupiter rocket. The Jupiter vehicle was a direct derivative of the Redstone. The Army Ballistic Missile Agency (ABMA) at Redstone Arsenal, Alabama, continued Jupiter development into a successful intermediate ballistic missile, even though the Department of Defense directed its operational development to the Air Force. ABMA maintained a role in Jupiter RD, including high-altitude launches that added to ABMA's understanding of rocket vehicle operations in the near-Earth space environment. It was knowledge that paid handsome dividends later.

  9. Early Rockets

    NASA Image and Video Library

    1957-03-01

    The Jupiter rocket was designed and developed by the Army Ballistic Missile Agency (ABMA). ABMA launched the Jupiter-A at Cape Canaveral, Florida, on March 1, 1957. The Jupiter vehicle was a direct derivative of the Redstone. The Army Ballistic Missile Agency (ABMA) at Redstone Arsenal, Alabama, continued Jupiter development into a successful intermediate ballistic missile, even though the Department of Defense directed its operational development to the Air Force. ABMA maintained a role in Jupiter RD, including high-altitude launches that added to ABMA's understanding of rocket vehicle operations in the near-Earth space environment. It was knowledge that paid handsome dividends later.

  10. Air-Powered Rockets.

    ERIC Educational Resources Information Center

    Rodriguez, Charley; Raynovic, Jim

    This document describes methods for designing and building two types of rockets--rockets from paper and rockets from bottles. Devices used for measuring the heights that the rockets obtain are also discussed. (KHR)

  11. Sounding rockets explore the ionosphere

    SciTech Connect

    Mendillo, M. )

    1990-08-01

    It is suggested that small, expendable, solid-fuel rockets used to explore ionospheric plasma can offer insight into all the processes and complexities common to space plasma. NASA's sounding rocket program for ionospheric research focuses on the flight of instruments to measure parameters governing the natural state of the ionosphere. Parameters include input functions, such as photons, particles, and composition of the neutral atmosphere; resultant structures, such as electron and ion densities, temperatures and drifts; and emerging signals such as photons and electric and magnetic fields. Systematic study of the aurora is also conducted by these rockets, allowing sampling at relatively high spatial and temporal rates as well as investigation of parameters, such as energetic particle fluxes, not accessible to ground based systems. Recent active experiments in the ionosphere are discussed, and future sounding rocket missions are cited.

  12. Advanced rocket propulsion

    NASA Technical Reports Server (NTRS)

    Obrien, Charles J.

    1993-01-01

    Existing NASA research contracts are supporting development of advanced reinforced polymer and metal matrix composites for use in liquid rocket engines of the future. Advanced rocket propulsion concepts, such as modular platelet engines, dual-fuel dual-expander engines, and variable mixture ratio engines, require advanced materials and structures to reduce overall vehicle weight as well as address specific propulsion system problems related to elevated operating temperatures, new engine components, and unique operating processes. High performance propulsion systems with improved manufacturability and maintainability are needed for single stage to orbit vehicles and other high performance mission applications. One way to satisfy these needs is to develop a small engine which can be clustered in modules to provide required levels of total thrust. This approach should reduce development schedule and cost requirements by lowering hardware lead times and permitting the use of existing test facilities. Modular engines should also reduce operational costs associated with maintenance and parts inventories.

  13. Regression rate behaviors of HTPB-based propellant combinations for hybrid rocket motor

    NASA Astrophysics Data System (ADS)

    Sun, Xingliang; Tian, Hui; Li, Yuelong; Yu, Nanjia; Cai, Guobiao

    2016-02-01

    The purpose of this paper is to characterize the regression rate behavior of hybrid rocket motor propellant combinations, using hydrogen peroxide (HP), gaseous oxygen (GOX), nitrous oxide (N2O) as the oxidizer and hydroxyl-terminated poly-butadiene (HTPB) as the based fuel. In order to complete this research by experiment and simulation, a hybrid rocket motor test system and a numerical simulation model are established. Series of hybrid rocket motor firing tests are conducted burning different propellant combinations, and several of those are used as references for numerical simulations. The numerical simulation model is developed by combining the Navies-Stokes equations with the turbulence model, one-step global reaction model, and solid-gas coupling model. The distribution of regression rate along the axis is determined by applying simulation mode to predict the combustion process and heat transfer inside the hybrid rocket motor. The time-space averaged regression rate has a good agreement between the numerical value and experimental data. The results indicate that the N2O/HTPB and GOX/HTPB propellant combinations have a higher regression rate, since the enhancement effect of latter is significant due to its higher flame temperature. Furthermore, the containing of aluminum (Al) and/or ammonium perchlorate(AP) in the grain does enhance the regression rate, mainly due to the more energy released inside the chamber and heat feedback to the grain surface by the aluminum combustion.

  14. Early Rockets

    NASA Image and Video Library

    2004-04-15

    By the end of the 19th Century, a Russian theorist, Konstantian Tsiolkovsky, was examining the fundamental scientific theories behind rocketry. He made some pioneering studies in liquid chemical rocket concepts and recommended liquid oxygen and liquid hydrogen as the optimum propellants. In the 1920's, Tsiolkovsky analyzed and mathematically formulated the technique for staged vehicles to reach escape velocities from Earth.

  15. Early Rockets

    NASA Image and Video Library

    1990-07-25

    An Atlas Centaur rocket (AC-S9) was launched from Cape Canaveral Air Force Station complex 36B carrying into orbit the Combined Release and Radiation Effects Satellite (CRRES) spacecraft. CRRES was a joint NASA/Air Force mission to study the effects of chemical release on the Earth’s atmosphere and magnetosphere.

  16. Early Rockets

    NASA Image and Video Library

    1959-05-28

    On May 28, 1959, a Jupiter Intermediate Range Ballistic Missile provided by a U.S. Army team in Redstone Arsenal, Alabama, launched a nose cone carrying Baker, A South American squirrel monkey and Able, An American-born rhesus monkey. This photograph shows Able after recovery of the nose cone of the Jupiter rocket by U.S.S. Kiowa.

  17. Early Rockets

    NASA Image and Video Library

    1959-08-14

    The Juno II launch vehicle, shown here, was a modified Jupiter Intermediate-Range Ballistic missionile, developed by Dr. Wernher von Braun and the rocket team at Redstone Arsenal in Huntsville, Alabama. Between December 1958 and April 1961, the Juno II launched space probes Pioneer III and IV, as well as Explorer satellites VII, VIII and XI.

  18. Early Rockets

    NASA Image and Video Library

    1940-01-01

    In this undated file photo, probably from World War II, a V-2 rocket emerges from its camouflaged shelter. The team of German engineers and scientists who developed the V-2 came to the United States after World War II and worked for the U. S. Army at Fort Bliss, Texas and Redstone Arsenal in Huntsville, Alabama.

  19. Early Rockets

    NASA Image and Video Library

    1944-01-01

    German technicians stack the various stages of the V-2 rocket in this undated photograph. The team of German engineers and scientists who developed the V-2 came to the United States at the end of World War II and worked for the U. S. Army at Fort Bliss, Texas, and Redstone Arsenal in Huntsville, Alabama.

  20. Saving Lives With Rocket Power

    NASA Technical Reports Server (NTRS)

    2000-01-01

    Thiokol Propulsion uses NASA's surplus rocket fuel to produce a flare that can safely destroy land mines. Through a Memorandum of Agreement between Thiokol and Marshall Space Flight Center, Thiokol uses the scrap Reusable Solid Rocket Motor (RSRM) propellant. The resulting Demining Device was developed by Thiokol with the help of DE Technologies. The Demining Device neutralizes land mines in the field without setting them off. The Demining Device flare is placed next to an uncovered land mine. Using a battery-triggered electric match, the flare is then ignited. Using the excess and now solidified rocket fuel, the flare burns a hole in the mine's case and ignites the explosive contents. Once the explosive material is burned away, the mine is disarmed and no longer dangerous.

  1. Reforming of fuel inside fuel cell generator

    DOEpatents

    Grimble, Ralph E.

    1988-01-01

    Disclosed is an improved method of reforming a gaseous reformable fuel within a solid oxide fuel cell generator, wherein the solid oxide fuel cell generator has a plurality of individual fuel cells in a refractory container, the fuel cells generating a partially spent fuel stream and a partially spent oxidant stream. The partially spent fuel stream is divided into two streams, spent fuel stream I and spent fuel stream II. Spent fuel stream I is burned with the partially spent oxidant stream inside the refractory container to produce an exhaust stream. The exhaust stream is divided into two streams, exhaust stream I and exhaust stream II, and exhaust stream I is vented. Exhaust stream II is mixed with spent fuel stream II to form a recycle stream. The recycle stream is mixed with the gaseous reformable fuel within the refractory container to form a fuel stream which is supplied to the fuel cells. Also disclosed is an improved apparatus which permits the reforming of a reformable gaseous fuel within such a solid oxide fuel cell generator. The apparatus comprises a mixing chamber within the refractory container, means for diverting a portion of the partially spent fuel stream to the mixing chamber, means for diverting a portion of exhaust gas to the mixing chamber where it is mixed with the portion of the partially spent fuel stream to form a recycle stream, means for injecting the reformable gaseous fuel into the recycle stream, and means for circulating the recycle stream back to the fuel cells.

  2. Reforming of fuel inside fuel cell generator

    DOEpatents

    Grimble, R.E.

    1988-03-08

    Disclosed is an improved method of reforming a gaseous reformable fuel within a solid oxide fuel cell generator, wherein the solid oxide fuel cell generator has a plurality of individual fuel cells in a refractory container, the fuel cells generating a partially spent fuel stream and a partially spent oxidant stream. The partially spent fuel stream is divided into two streams, spent fuel stream 1 and spent fuel stream 2. Spent fuel stream 1 is burned with the partially spent oxidant stream inside the refractory container to produce an exhaust stream. The exhaust stream is divided into two streams, exhaust stream 1 and exhaust stream 2, and exhaust stream 1 is vented. Exhaust stream 2 is mixed with spent fuel stream 2 to form a recycle stream. The recycle stream is mixed with the gaseous reformable fuel within the refractory container to form a fuel stream which is supplied to the fuel cells. Also disclosed is an improved apparatus which permits the reforming of a reformable gaseous fuel within such a solid oxide fuel cell generator. The apparatus comprises a mixing chamber within the refractory container, means for diverting a portion of the partially spent fuel stream to the mixing chamber, means for diverting a portion of exhaust gas to the mixing chamber where it is mixed with the portion of the partially spent fuel stream to form a recycle stream, means for injecting the reformable gaseous fuel into the recycle stream, and means for circulating the recycle stream back to the fuel cells. 1 fig.

  3. Method for removing acid gases from a gaseous stream

    DOEpatents

    Gorin, Everett; Zielke, Clyde W.

    1981-01-01

    In a process for hydrocracking a heavy aromatic polynuclear carbonaceous feedstock containing reactive alkaline constituents to produce liquid hydrocarbon fuels boiling below about 475.degree. C. at atmospheric pressure by contacting the feedstock with hydrogen in the presence of a molten metal halide catalyst, thereafter separating a gaseous stream containing hydrogen, at least a portion of the hydrocarbon fuels and acid gases from the molten metal halide and regenerating the molten metal halide, thereby producing a purified molten metal halide stream for recycle to the hydrocracking zone, an improvement comprising; contacting the gaseous acid gas, hydrogen and hydrocarbon fuels-containing stream with the feedstock containing reactive alkaline constituents to remove acid gases from the acid gas containing stream. Optionally at least a portion of the hydrocarbon fuels are separated from gaseous stream containing hydrogen, hydrocarbon fuels and acid gases prior to contacting the gaseous stream with the feedstock.

  4. The beginnings. [Of Nuclear Engine for Rocket Vehicles Application

    SciTech Connect

    Bohl, R.J.; Kirk, W.L.; Holman, R.R.; Westinghouse Electric Corp., Pittsburgh, PA )

    1989-06-01

    The development of the nuclear rocket engine called NERVA (Nuclear Engine for Rocket Vehicle Application) is described. The choice of fuel element, required rocket parameters, NERVA project objectives, division of responsibilities among different organizations, and NERVA design configuration are reviewed. Progress that has been made in the development of NERVA is addressed.

  5. Mars Rocket Propulsion System

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert; Harber, Dan; Nabors, Sammy

    2008-01-01

    A report discusses the methane and carbon monoxide/LOX (McLOx) rocket for ascent from Mars as well as other critical space propulsion tasks. The system offers a specific impulse over 370 s roughly 50 s higher than existing space-storable bio-propellants. Current Mars in-situ propellant production (ISPP) technologies produce impure methane and carbon monoxide in various combinations. While separation and purification of methane fuel is possible, it adds complexity to the propellant production process and discards an otherwise useful fuel product. The McLOx makes such complex and wasteful processes unnecessary by burning the methane/CO mixtures produced by the Mars ISPP systems without the need for further refinement. Despite the decrease in rocket-specific impulse caused by the CO admixture, the improvement offered by concomitant increased propellant density can provide a net improvement in stage performance. One advantage is the increase of the total amount of propellant produced, but with a decrease in mass and complexity of the required ISPP plant. Methane/CO fuel mixtures also may be produced by reprocessing the organic wastes of a Moon base or a space station, making McLOx engines key for a human Lunar initiative or the International Space Station (ISS) program. Because McLOx propellant components store at a common temperature, very lightweight and compact common bulkhead tanks can be employed, improving overall stage performance further.

  6. The four INTA-300 rocket prototypes

    NASA Astrophysics Data System (ADS)

    Calero, J. S.

    1985-03-01

    A development history and performance capability assessment is presented for the INTA-300 'Flamenco' sounding rocket prototype specimens. The Flamenco is a two-stage solid fuel rocket, based on British sounding rocket technology, that can lift 50 km payloads to altitudes of about 300 km. The flight of the first two prototypes, in 1974 and 1975, pointed to vibration problems which reduced the achievable apogee, and the third prototype's flight was marred by a premature detonation that destroyed the rocket. The fourth Flamenco flight, however, yielded much reliable data.

  7. ION ROCKET ENGINE

    DOEpatents

    Ehlers, K.W.; Voelker, F. III

    1961-12-19

    A thrust generating engine utilizing cesium vapor as the propellant fuel is designed. The cesium is vaporized by heat and is passed through a heated porous tungsten electrode whereby each cesium atom is fonized. Upon emergfng from the tungsten electrode, the ions are accelerated rearwardly from the rocket through an electric field between the tungsten electrode and an adjacent accelerating electrode grid structure. To avoid creating a large negative charge on the space craft as a result of the expulsion of the positive ions, a source of electrons is disposed adjacent the ion stream to neutralize the cesium atoms following acceleration thereof. (AEC)

  8. Rocket Noise Prediction Program

    NASA Technical Reports Server (NTRS)

    Margasahayam, Ravi; Caimi, Raoul

    1999-01-01

    A comprehensive, automated, and user-friendly software program was developed to predict the noise and ignition over-pressure environment generated during the launch of a rocket. The software allows for interactive modification of various parameters affecting the generated noise environment. Predictions can be made for different launch scenarios and a variety of vehicle and launch mount configurations. Moreover, predictions can be made for both near-field and far-field locations on the ground and any position on the vehicle. Multiple engine and fuel combinations can be addressed, and duct geometry can be incorporated efficiently. Applications in structural design are addressed.

  9. Rocket plume flowfield characterization using laser Rayleigh scattering

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J.; Weiss, Jonathan M.

    1992-01-01

    A Doppler-resolved laser Rayleigh scattering diagnostic was applied to a 111 N thrust, regenerative and fuel-film cooled, gaseous hydrogen/gaseous oxygen rocket engine. The axial and radial mean gas velocities were measured from the net Doppler shifts observed for two different scattering angles. Translational temperatures and number densities were estimated from the Doppler widths and scattered intensities, respectively, by assuming that water was the dominant scattering species in the exhaust. The experimental results are compared with theoretical predictions from a full Navier-Stokes code (RD/RPLUS) and the JANNAF Two-Dimensional Kinetics (TDK) and Standardized Plume Flowfield (SPF-II) codes. Discrepancies between the measured and predicted axial velocities, temperatures, and number densities are evident. Radial velocity measurements, however, show excellent agreement with predictions. The discrepancies are attributed primarily to inefficient mixing and combustion caused by the injection of excessive oxidizer along one side of the thrust chamber. Thrust and mass flow rate estimates obtained from the Rayleigh measurements show excellent agreement with the globally measured values.

  10. Rocket plume flowfield characterization using laser Rayleigh scattering

    NASA Astrophysics Data System (ADS)

    Zupanc, Frank J.; Weiss, Jonathan M.

    1992-07-01

    A Doppler-resolved laser Rayleigh scattering diagnostic was applied to a 111 N thrust, regenerative and fuel-film cooled, gaseous hydrogen/gaseous oxygen rocket engine. The axial and radial mean gas velocities were measured from the net Doppler shifts observed for two different scattering angles. Translational temperatures and number densities were estimated from the Doppler widths and scattered intensities, respectively, by assuming that water was the dominant scattering species in the exhaust. The experimental results are compared with theoretical predictions from a full Navier-Stokes code (RD/RPLUS) and the JANNAF Two-Dimensional Kinetics (TDK) and Standardized Plume Flowfield (SPF-II) codes. Discrepancies between the measured and predicted axial velocities, temperatures, and number densities are evident. Radial velocity measurements, however, show excellent agreement with predictions. The discrepancies are attributed primarily to inefficient mixing and combustion caused by the injection of excessive oxidizer along one side of the thrust chamber. Thrust and mass flow rate estimates obtained from the Rayleigh measurements show excellent agreement with the globally measured values.

  11. Rocket plume flowfield characterization using laser Rayleigh scattering

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J.; Weiss, Jonathan M.

    1992-01-01

    A Doppler-resolved laser Rayleigh scattering diagnostic was applied to a 111 N thrust, regenerative and fuel-film cooled, gaseous hydrogen/gaseous oxygen rocket engine. The axial and radial mean gas velocities were measured from the net Doppler shifts observed for two different scattering angles. Translational temperatures and number densities were estimated from the Doppler widths and scattered intensities, respectively, by assuming that water was the dominant scattering species in the exhaust. The experimental results are compared with theoretical predictions from a full Navier-Stokes code (RD/RPLUS) and the JANNAF Two-Dimensional Kinetics (TDK) and Standardized Plume Flowfield (SPF-II) codes. Discrepancies between the measured and predicted axial velocities, temperatures, and number densities are evident. Radial velocity measurements, however, show excellent agreement with predictions. The discrepancies are attributed primarily to inefficient mixing and combustion caused by the injection of excessive oxidizer along one side of the thrust chamber. Thrust and mass flow rate estimates obtained from the Rayleigh measurements show excellent agreement with the globally measured values.

  12. A new facility for advanced rocket propulsion research

    NASA Technical Reports Server (NTRS)

    Zoeckler, Joseph G.; Green, James M.; Raitano, Paul

    1993-01-01

    A new test facility was constructed at the NASA Lewis Research Center Rocket Laboratory for the purpose of conducting rocket propulsion research at up to 8.9 kN (2000 lbf) thrust, using liquid oxygen and gaseous hydrogen propellants. A laser room adjacent to the test cell provides access to the rocket engine for advanced laser diagnostic systems. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods, with rapid turnover between programs. These capabilities make the new test facility an important asset for basic and applied rocket propulsion research.

  13. Devising rocket power for smaller engines

    SciTech Connect

    Burruss, R.

    1996-04-01

    Compact, high-power engines that burn fuel and oxygen could be made by winding copper tubing in a helix around boiler sections. With more than 1,000 horsepower per pound of engine weight, liquid-fueled rockets have the highest specific power of any engines designed for sustained operation. Yet those engines generally run for about only 1,000 seconds--nowhere near the sustained operation time for lower-power automotive and aircraft engines of more than 1,000 hours. In theory, at least, a fuel/oxygen rocket can be built that combines the best of both classes: high specific power (from perhaps two to 10 times that of a gas turbine) and a 1,000-hour service life. Such an engine would almost certainly be possible if the rocket`s exhaust gases could be simultaneously cooled and expanded by mixing water with the rocket`s exhaust and boiling it before it reaches the turbine. The technology itself is not new. variations of these rocket-turbine-type engines, for example, powered torpedoes during World War I. Some 30 years later, German V-2 rockets used fuel pumps, driven by the reaction of hydrogen peroxide with hydrocarbon fuels, to produce high-pressure steam that was directed against a turbine. Alternatively, fuel/oxygen combustion could produce steam to drive a piston engine. Either way, the challenge remains to construct a compact, long-service-life, high-specific-power boiler that burns fuel and oxygen. The new type of engine could be derived from recent research on electric vehicles (EVs).

  14. Early Rockets

    NASA Image and Video Library

    2004-04-15

    This photograph is of the engine for the Redstone rocket. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the "reliable workhorse" for America's early space program. As an example of its versatility, the Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile.

  15. Early Rockets

    NASA Image and Video Library

    2004-04-15

    According to one ancient legend, a Chinese official named Wan Hoo attempted a flight to the moon using a large wicker chair to which were fastened 47 large rockets. Forty seven assistants, each armed with a torch, rushed forward to light the fuses. In a moment, there was a tremendous roar accompanied by billowing clouds of smoke. When the smoke cleared, the flying chair and Wan Hoo were gone.

  16. Rocket Tablet,

    DTIC Science & Technology

    1984-09-12

    is a vast and desolate world, this is a strip of mir- aculous land! How many struggling dramas full of power and * grandeur were cheered, resisted and...rocket officers and men, a group enormous and powerful , marched into this land soaked with the fresh blood of our ancestors. This place is about to...and tough pestering said he wanted an American aircraft ob- tained on the battlefield to transport goods from Lanzhou, Xian, Beijing, Guangzhou and

  17. Early Rockets

    NASA Image and Video Library

    1958-05-28

    On May 28, 1958, Jupiter Intermediate Range Ballistic Missile provided by U.S. Army team in Huntsville, Alabama, launched a nose cone carrying Baker, a South American squirrel monkey and Able, an American-born rhesus monkey. Baker, pictured here and commonly known as "Miss Baker", was later given a home at the U.S. Space and Rocket Center until her death on November 29, 1984. Able died in 1958. (Photo - Courtesy of Huntsville/Madison County Public Library)

  18. Early Rockets

    NASA Image and Video Library

    1957-10-04

    The Army Ballistic Missile Agency incorporated the von Braun team in key positions with Dr. von Braun as a head of the Development Operations Division. On October 4, 1957, the Nation was shocked when the Russians launched Sputnik, the world's first artificial satellite. Two months later, the United States suffered disappointment when a Navy Vanguard rocket, with its satellite payload, failed to develop sufficient thrust and toppled over on the launch pad.

  19. Two-dimensional motions of rockets

    NASA Astrophysics Data System (ADS)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

    We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the descending parts of the trajectories tend to be gentler and straighter slopes than the ascending parts for relatively large launching angles due to the non-vanishing thrusts. We discuss the ranges, the maximum altitudes and the engine performances of the rockets. It seems that the exponential fuel exhaustion can be the most potent engine for the longest and highest flights.

  20. Technique for predicting high-frequency stability characteristics of gaseous-propellant combustors

    NASA Technical Reports Server (NTRS)

    Priem, R. J.; Jefferson, Y. S. Y.

    1973-01-01

    A technique for predicting the stability characteristics of a gaseous-propellant rocket combustion system is developed based on a model that assumes coupling between the flow through the injector and the oscillating chamber pressure. The theoretical model uses a lumped parameter approach for the flow elements in the injection system plus wave dynamics in the combustion chamber. The injector flow oscillations are coupled to the chamber pressure oscillations with a delay time. Frequency and decay (or growth) rates are calculated for various combustor design and operating parameters to demonstrate the influence of various parameters on stability. Changes in oxidizer design parameters had a much larger influence on stability than a similar change in fuel parameters. A complete description of the computer program used to make these calculations is given in an appendix.

  1. Numerical and Experimental Study of Mixing Properties of Gaseous Fuels Jets Including Hydrogen and Methane Into the non-Swirl Main Flow in a Premixer Configuration

    NASA Astrophysics Data System (ADS)

    Akbari, Amin

    The mixing of fuel and air has a significant impact on overall operation efficiency and emissions performance of combustion systems, especially in lean combustion applications. As a result, developing an understanding of the processes associated with the fuel/air mixing is important. In parallel with the evolution of lean combustion, a new generation of fuels is emerging as an alternative to conventional fuels. Thus, it is desirable to study the mixing properties of different fuels from conventional resources, such as methane, as well as from renewable resources, such as hydrogen. One tool that is available to study mixing in complex (e.g., turbulent and elliptic) flows is computational fluid dynamics (CFD). In the present work, mixing of hydrogen and methane into air, for example, is simulated using various CFD approaches. Fuel is injected either co-flowing to the air flow ("axial injection") or perpendicular to the air flow ("radial injection"). The quality of the simulations is evaluated by comparing the numerical results with experimental measurements. Qualitative and quantitative comparisons are used to evaluate the relative accuracy of different CFD approaches to simulate the mixing characteristics. Reynolds Averaged Navier-Stokes (RANS) turbulent models are utilized to model all the cases as steady turbulent models. Moreover, unsteady turbulent models, such as Unsteady RANS, and Large Eddy Simulation (LES) are used to provide information about unsteady features in selected cases. The sensitivity of numerical predictions to different RANS turbulence models as well as to different turbulent Schmidt numbers are explored. The results indicate more sensitivity to turbulence models for radial injection configurations. However, for the axial configuration, more sensitivity to Sct is observed. In general, the RSM turbulence model with Sc t=0.7 provides the most promising predictions for various combination of different fuels and injection types.

  2. Laser diagnostics for small rockets

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J.; Degroot, Wilhelmus A.

    1993-01-01

    Two nonintrusive flowfield diagnostics based on spectrally-resolved elastic (Rayleigh) and inelastic (Raman) laser light scattering were developed for obtaining local flowfield measurements in low-thrust gaseous H2/O2 rocket engines. The objective is to provide an improved understanding of phenomena occurring in small chemical rockets in order to facilitate the development and validation of advanced computational fluid dynamics (CFD) models for analyzing engine performance. The laser Raman scattering diagnostic was developed to measure major polyatomic species number densities and rotational temperatures in the high-density flowfield region extending from the injector through the chamber throat. Initial application of the Raman scattering diagnostic provided O2 number density and rotational temperature measurements in the exit plane of a low area-ratio nozzle and in the combustion chamber of a two-dimensional, optically-accessible rocket engine. In the low-density nozzle exit plane region where the Raman signal is too weak, a Doppler-resolved laser Rayleigh scattering diagnostic was developed to obtain axial and radial mean gas velocities, and in certain cases, H2O translational temperature and number density. The results from these measurements were compared with theoretical predictions from the RPLUS CFD code for analyzing rocket engine performance. Initial conclusions indicate that a detailed and rigorous modeling of the injector is required in order to make direct comparisons between laser diagnostic measurements and CFD predictions at the local level.

  3. 14 CFR 34.71 - Compliance with gaseous emission standards.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... TRANSPORTATION AIRCRAFT FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE ENGINE POWERED AIRPLANES Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft and Aircraft Gas Turbine Engines) § 34.71...

  4. Separation of gaseous hydrogen from a water-hydrogen mixture in a fuel cell power system operating in a weightless environment

    NASA Technical Reports Server (NTRS)

    Romanowski, William E. (Inventor); Suljak, George T. (Inventor)

    1989-01-01

    A fuel cell power system for use in a weightless environment, such as in space, includes a device for removing water from a water-hydrogen mixture condensed from the exhaust from the fuel cell power section of the system. Water is removed from the mixture in a centrifugal separator, and is fed into a holding, pressure operated water discharge valve via a Pitot tube. Entrained nondissolved hydrogen is removed from the Pitot tube by a bleed orifice in the Pitot tube before the water reaches the water discharge valve. Water discharged from the valve thus has a substantially reduced hydrogen content.

  5. Development and Application of a Sample Holder for In Situ Gaseous TEM Studies of Membrane Electrode Assemblies for Polymer Electrolyte Fuel Cells.

    PubMed

    Kamino, Takeo; Yaguchi, Toshie; Shimizu, Takahiro

    2017-10-01

    Polymer electrolyte fuel cells hold great potential for stationary and mobile applications due to high power density and low operating temperature. However, the structural changes during electrochemical reactions are not well understood. In this article, we detail the development of the sample holder equipped with gas injectors and electric conductors and its application to a membrane electrode assembly of a polymer electrolyte fuel cell. Hydrogen and oxygen gases were simultaneously sprayed on the surfaces of the anode and cathode catalysts of the membrane electrode assembly sample, respectively, and observation of the structural changes in the catalysts were simultaneously carried out along with measurement of the generated voltages.

  6. Parametric study and performance analysis of hybrid rocket motors with double-tube configuration

    NASA Astrophysics Data System (ADS)

    Yu, Nanjia; Zhao, Bo; Lorente, Arnau Pons; Wang, Jue

    2017-03-01

    The practical implementation of hybrid rocket motors has historically been hampered by the slow regression rate of the solid fuel. In recent years, the research on advanced injector designs has achieved notable results in the enhancement of the regression rate and combustion efficiency of hybrid rockets. Following this path, this work studies a new configuration called double-tube characterized by injecting the gaseous oxidizer through a head end injector and an inner tube with injector holes distributed along the motor longitudinal axis. This design has demonstrated a significant potential for improving the performance of hybrid rockets by means of a better mixing of the species achieved through a customized injection of the oxidizer. Indeed, the CFD analysis of the double-tube configuration has revealed that this design may increase the regression rate over 50% with respect to the same motor with a conventional axial showerhead injector. However, in order to fully exploit the advantages of the double-tube concept, it is necessary to acquire a deeper understanding of the influence of the different design parameters in the overall performance. In this way, a parametric study is carried out taking into account the variation of the oxidizer mass flux rate, the ratio of oxidizer mass flow rate injected through the inner tube to the total oxidizer mass flow rate, and injection angle. The data for the analysis have been gathered from a large series of three-dimensional numerical simulations that considered the changes in the design parameters. The propellant combination adopted consists of gaseous oxygen as oxidizer and high-density polyethylene as solid fuel. Furthermore, the numerical model comprises Navier-Stokes equations, k-ε turbulence model, eddy-dissipation combustion model and solid-fuel pyrolysis, which is computed through user-defined functions. This numerical model was previously validated by analyzing the computational and experimental results obtained for

  7. 14 CFR 34.64 - Sampling and analytical procedures for measuring gaseous exhaust emissions.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE... Turbine Engines) § 34.64 Sampling and analytical procedures for measuring gaseous exhaust emissions....

  8. 14 CFR 34.64 - Sampling and analytical procedures for measuring gaseous exhaust emissions.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE... Turbine Engines) § 34.64 Sampling and analytical procedures for measuring gaseous exhaust emissions....

  9. Evaluation by Rocket Combustor of C/C Composite Cooled Structure Using Metallic Cooling Tubes

    NASA Astrophysics Data System (ADS)

    Takegoshi, Masao; Ono, Fumiei; Ueda, Shuichi; Saito, Toshihito; Hayasaka, Osamu

    In this study, the cooling performance of a C/C composite material structure with metallic cooling tubes fixed by elastic force without chemical bonding was evaluated experimentally using combustion gas in a rocket combustor. The C/C composite chamber was covered by a stainless steel outer shell to maintain its airtightness. Gaseous hydrogen as a fuel and gaseous oxygen as an oxidizer were used for the heating test. The surface of these C/C composites was maintained below 1500 K when the combustion gas temperature was about 2800 K and the heat flux to the combustion chamber wall was about 9 MW/m2. No thermal damage was observed on the stainless steel tubes that were in contact with the C/C composite materials. The results of the heating test showed that such a metallic tube-cooled C/C composite structure is able to control the surface temperature as a cooling structure (also as a heat exchanger) as well as indicated the possibility of reducing the amount of coolant even if the thermal load to the engine is high. Thus, application of this metallic tube-cooled C/C composite structure to reusable engines such as a rocket-ramjet combined-cycle engine is expected.

  10. Theoretical performance of liquid ammonia, hydrazine and mixture of liquid ammonia and hydrazine as fuels with liquid oxygen biflouride as oxidant for rocket engines : I-mixture of liquid ammonia and hydrazine

    NASA Technical Reports Server (NTRS)

    Huff, Vearl N; Gordon, Sanford

    1952-01-01

    Theoretical performance for mixture of 36.3 percent liquid ammonia and 63.7 percent hydrazine with liquid oxygen bifluoride as rocket propellant was calculated on assumption of equilibrium composition during expansion for a wide range of fuel-oxidant and expansios ratios. Parameters included were specific impulse, combustion-chamber temperature, nozzle exit temperature, composition mean molecular weight, characteristic velocity, coefficient of thrust and ratio of nozzle-exit area to throat area. For chamber pressure of 300 pounds per square inch absolute and expansion to 1 atmosphere, maximum specific impulse was 295.8 pound-seconds per pound. Five percent by weight of water in the hydrazine lowered specific impulse from about one to three units over a wide range of weight-percent fuel.

  11. Gaseous and Particulate Emissions from Diesel Engines at Idle and under Load: Comparison of Biodiesel Blend and Ultralow Sulfur Diesel Fuels

    PubMed Central

    Chin, Jo-Yu; Batterman, Stuart A.; Northrop, William F.; Bohac, Stanislav V.; Assanis, Dennis N.

    2015-01-01

    Diesel exhaust emissions have been reported for a number of engine operating strategies, after-treatment technologies, and fuels. However, information is limited regarding emissions of many pollutants during idling and when biodiesel fuels are used. This study investigates regulated and unregulated emissions from both light-duty passenger car (1.7 L) and medium-duty (6.4 L) diesel engines at idle and load and compares a biodiesel blend (B20) to conventional ultralow sulfur diesel (ULSD) fuel. Exhaust aftertreatment devices included a diesel oxidation catalyst (DOC) and a diesel particle filter (DPF). For the 1.7 L engine under load without a DOC, B20 reduced brake-specific emissions of particulate matter (PM), elemental carbon (EC), nonmethane hydrocarbons (NMHCs), and most volatile organic compounds (VOCs) compared to ULSD; however, formaldehyde brake-specific emissions increased. With a DOC and high load, B20 increased brake-specific emissions of NMHC, nitrogen oxides (NOx), formaldehyde, naphthalene, and several other VOCs. For the 6.4 L engine under load, B20 reduced brake-specific emissions of PM2.5, EC, formaldehyde, and most VOCs; however, NOx brake-specific emissions increased. When idling, the effects of fuel type were different: B20 increased NMHC, PM2.5, EC, formaldehyde, benzene, and other VOC emission rates from both engines, and changes were sometimes large, e.g., PM2.5 increased by 60% for the 6.4 L/2004 calibration engine, and benzene by 40% for the 1.7 L engine with the DOC, possibly reflecting incomplete combustion and unburned fuel. Diesel exhaust emissions depended on the fuel type and engine load (idle versus loaded). The higher emissions found when using B20 are especially important given the recent attention to exposures from idling vehicles and the health significance of PM2.5. The emission profiles demonstrate the effects of fuel type, engine calibration, and emission control system, and they can be used as source profiles for apportionment

  12. Gaseous and Particulate Emissions from Diesel Engines at Idle and under Load: Comparison of Biodiesel Blend and Ultralow Sulfur Diesel Fuels.

    PubMed

    Chin, Jo-Yu; Batterman, Stuart A; Northrop, William F; Bohac, Stanislav V; Assanis, Dennis N

    2012-11-15

    Diesel exhaust emissions have been reported for a number of engine operating strategies, after-treatment technologies, and fuels. However, information is limited regarding emissions of many pollutants during idling and when biodiesel fuels are used. This study investigates regulated and unregulated emissions from both light-duty passenger car (1.7 L) and medium-duty (6.4 L) diesel engines at idle and load and compares a biodiesel blend (B20) to conventional ultralow sulfur diesel (ULSD) fuel. Exhaust aftertreatment devices included a diesel oxidation catalyst (DOC) and a diesel particle filter (DPF). For the 1.7 L engine under load without a DOC, B20 reduced brake-specific emissions of particulate matter (PM), elemental carbon (EC), nonmethane hydrocarbons (NMHCs), and most volatile organic compounds (VOCs) compared to ULSD; however, formaldehyde brake-specific emissions increased. With a DOC and high load, B20 increased brake-specific emissions of NMHC, nitrogen oxides (NOx), formaldehyde, naphthalene, and several other VOCs. For the 6.4 L engine under load, B20 reduced brake-specific emissions of PM2.5, EC, formaldehyde, and most VOCs; however, NOx brake-specific emissions increased. When idling, the effects of fuel type were different: B20 increased NMHC, PM2.5, EC, formaldehyde, benzene, and other VOC emission rates from both engines, and changes were sometimes large, e.g., PM2.5 increased by 60% for the 6.4 L/2004 calibration engine, and benzene by 40% for the 1.7 L engine with the DOC, possibly reflecting incomplete combustion and unburned fuel. Diesel exhaust emissions depended on the fuel type and engine load (idle versus loaded). The higher emissions found when using B20 are especially important given the recent attention to exposures from idling vehicles and the health significance of PM2.5. The emission profiles demonstrate the effects of fuel type, engine calibration, and emission control system, and they can be used as source profiles for apportionment

  13. Early Rockets

    NASA Image and Video Library

    1955-09-01

    Launch of a three-stage Vanguard (SLV-7) from Cape Canaveral, Florida, September 18, 1959. Designated Vanguard III, the 100-pound satellite was used to study the magnetic field and radiation belt. In September 1955, the Department of Defense recommended and authorized the new program, known as Project Vanguard, to launch Vanguard booster to carry an upper atmosphere research satellite in orbit. The Vanguard vehicles were used in conjunction with later booster vehicle such as the Thor and Atlas, and the technique of gimbaled (movable) engines for directional control was adapted to other rockets.

  14. CAMUI Type Hybrid Rocket as Small Scale Ballistic Flight Testbed

    NASA Astrophysics Data System (ADS)

    Nagata, Harunori; Uematsu, Tsutomu; Ito, Kenichi

    The authors have been developing CAMUI (Cascaded Multistage Impinging-jet) type hybrid rockets, explosive-flee small rocket motors. This is to downsize the scale of suborbital flight experiments on space related technology development. A key idea is a new fuel grain design to increase gasification rates of a solid fuels. By the new fuel grain design, the combustion gas repeatedly impinges on fuel surfaces to hasten the heat transfer to the fuel. Suborbital flight experiments by sounding rockets provide variety of test beds to accumulate basic technologies common to the next step of space development in Japan. By using hybrid rockets one can take the cost advantage of small-scale rocket experiments. This cost advantage improves robustness of space technology development projects by dispersion of risk.

  15. Experiment/Analytical Characterization of the RBCC Rocket-Ejector Mode

    NASA Technical Reports Server (NTRS)

    Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.; West, J.; Turner, James E. (Technical Monitor)

    2000-01-01

    Experimental and complementary CFD results from the study of the rocket-ejector mode of a Rocket Based Combined Cycle (RBCC) engine are presented and discussed. The experiments involved systematic flowfield measurements in a two-dimensional, variable geometry rocket-ejector system. The rocket-ejector system utilizes a single two-dimensional, gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a thorough understanding of the rocket-ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions. Overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (oxygen, hydrogen, nitrogen and water vapor). The experimental results for both the direct-connect and sea-level static configurations are compared with CFD predictions of the flowfield.

  16. Current and Future Critical Issues in Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Navaz, Homayun K.; Dix, Jeff C.

    1998-01-01

    The objective of this research was to tackle several problems that are currently of great importance to NASA. In a liquid rocket engine several complex processes take place that are not thoroughly understood. Droplet evaporation, turbulence, finite rate chemistry, instability, and injection/atomization phenomena are some of the critical issues being encountered in a liquid rocket engine environment. Pulse Detonation Engines (PDE) performance, combustion chamber instability analysis, 60K motor flowfield pattern from hydrocarbon fuel combustion, and 3D flowfield analysis for the Combined Cycle engine were of special interest to NASA. During the summer of 1997, we made an attempt to generate computational results for all of the above problems and shed some light on understanding some of the complex physical phenomena. For this purpose, the Liquid Thrust Chamber Performance (LTCP) code, mainly designed for liquid rocket engine applications, was utilized. The following test cases were considered: (1) Characterization of a detonation wave in a Pulse Detonation Tube; (2) 60K Motor wall temperature studies; (3) Propagation of a pressure pulse in a combustion chamber (under single and two-phase flow conditions); (4) Transonic region flowfield analysis affected by viscous effects; (5) Exploring the viscous differences between a smooth and a corrugated wall; and (6) 3D thrust chamber flowfield analysis of the Combined Cycle engine. It was shown that the LTCP-2D and LTCP-3D codes are capable of solving complex and stiff conservation equations for gaseous and droplet phases in a very robust and efficient manner. These codes can be run on a workstation and personal computers (PC's).

  17. Small hydrogen/oxygen rocket flowfield behavior from heat flux measurements

    NASA Technical Reports Server (NTRS)

    Reed, Brian D.

    1993-01-01

    The mixing and heat transfer phenomena in small rocket flow fields with fuel film cooling is not well understood. An instrumented, water-cooled chamber with a gaseous hydrogen/gaseous oxygen injector was used to gather steady-state inner and outer wall temperature profiles. The chamber was tested at 414 kPa (60 psia) chamber pressure, from mixture ratios of 3.41 to 8.36. Sixty percent of the fuel was used for film cooling. These temperature profiles were used as boundary conditions in a finite element analysis program, MSC/NASTRAN, to calculate the local radial and axial heat fluxes in the chamber wall. The normal heat fluxes were then calculated and used as a diagnostic of the rocket's flow field behavior. The normal heat fluxes determined were on the order of 1.0 to 3.0 MW/meters squared (0.6 to 1.8 Btu/sec-inches squared). In the cases where mixture ratio was 5 or above, there was a sharp local heat flux maximum in the barrel section of the chamber. This local maximum seems to indicate a reduction or breakdown of the fuel film cooling layer, possibly due to increased mixing in the shear layer between the film and core flows. However, the flow was thought to be completely laminar, as the throat Reynolds numbers were below 50,000 for all the cases. The increased mixing in the shear layer in the higher mixture ratio cases appeared not to be due to the transition of the flow from laminar to turbulent, but rather due to increased reactions between the hydrogen film and oxidizer-rich core flows.

  18. Rocket Science at the Nanoscale.

    PubMed

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

    Autonomous propulsion at the nanoscale represents one of the most challenging and demanding goals in nanotechnology. Over the past decade, numerous important advances in nanotechnology and material science have contributed to the creation of powerful self-propelled micro/nanomotors. In particular, micro- and nanoscale rockets (MNRs) offer impressive capabilities, including remarkable speeds, large cargo-towing forces, precise motion controls, and dynamic self-assembly, which have paved the way for designing multifunctional and intelligent nanoscale machines. These multipurpose nanoscale shuttles can propel and function in complex real-life media, actively transporting and releasing therapeutic payloads and remediation agents for diverse biomedical and environmental applications. This review discusses the challenges of designing efficient MNRs and presents an overview of their propulsion behavior, fabrication methods, potential rocket fuels, navigation strategies, practical applications, and the future prospects of rocket science and technology at the nanoscale.

  19. Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Engine Calorimeter Heat Transfer Measurements and Analysis

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    1997-01-01

    A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. The analyses used the data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt %, 5-wt%, and 55-wt% loadings of aluminum with silicon dioxide gellant, and gaseous oxygen as the oxidizer. Heat transfer was computed based on measurements using calorimeter rocket chamber and nozzle hardware with a total of 31 cooling channels. A gelled fuel coating formed in the 0-, 5- and 55-wt% engines, and the coating was composed of unburned gelled fuel and partially combusted RP-1. The coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber for the 0- and 5-wt% RP-1/Al. This heat flux reduction effect was analyzed by comparing engine runs and the changes in the heat flux during a run as well as from run to run. Heat transfer and time-dependent heat flux analyses and interpretations are provided. The 5- and 55-wt% RP-1/Al fueled engines had the highest chamber heat fluxes, with the 5-wt% fuel having the highest throat flux. This result is counter to the predicted result, where the 55 wt% fuel has the highest combustion and throat temperature, and therefore implies that it would deliver the highest throat heat flux. The 5-wt% RP-1/Al produced the most influence on the engine heat transfer and the heat flux reduction was caused by the formation of a gelled propellant layer in the chamber and nozzle.

  20. Gaseous hydrogen/oxygen injector performance characterization

    NASA Technical Reports Server (NTRS)

    Degroot, W. A.; Tsuei, H. H.

    1994-01-01

    Results are presented of spontaneous Raman scattering measurements in the combustion chamber of a 110 N thrust class gaseous hydrogen/oxygen rocket. Temperature, oxygen number density, and water number density profiles at the injector exit plane are presented. These measurements are used as input profiles to a full Navier-Stokes computational fluid dynamics (CFD) code. Predictions of this code while using the measured profiles are compared with predictions while using assumed uniform injector profiles. Axial and radial velocity profiles derived from both sets of predictions are compared with Rayleigh scattering measurements in the exit plane of a 33:1 area ratio nozzle. Temperature and number density Raman scattering measurements at the exit plane of a test rocket with a 1:1.36 area ratio nozzle are also compared with results from both sets of predictions.

  1. Early Rockets

    NASA Image and Video Library

    2004-04-15

    Legendary characters used the power of mythology to fly through the heavens. About 200 BC, a Greek inventor known as Hero of Alexandria came up with a new invention that depended on the mechanical interaction of heat and water. He invented a rocket-like device called an aeolipile. It used steam for propulsion. Hero mounted a sphere on top of a water kettle. A fire below the kettle turned the water into steam, and the gas traveled through the pipes to the sphere. Two L-shaped tubes on opposite sides of the sphere allowed the gas to escape, and in doing so gave a thrust to the sphere that caused it to rotate.

  2. Early Rockets

    NASA Image and Video Library

    1958-05-15

    Redstone missile No. 1002 on the launch pad at Cape Canaveral, Florida, on May 16, 1958. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and production was begun in 1952. Redstone rockets became the "reliable workhorse" for America's early space program. As an example of the versatility, Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile

  3. Early Rockets

    NASA Image and Video Library

    2004-04-15

    The image depicts Redstone missile being erected. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the "reliable workhorse" for America's early space program. As an example of the versatility, Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile

  4. Results of Small-scale Solid Rocket Combustion Simulator testing at Marshall Space Flight Center

    NASA Astrophysics Data System (ADS)

    Goldberg, Benjamin E.; Cook, Jerry

    1993-06-01

    The Small-scale Solid Rocket Combustion Simulator (SSRCS) program was established at the Marshall Space Flight Center (MSFC), and used a government/industry team consisting of Hercules Aerospace Corporation, Aerotherm Corporation, United Technology Chemical Systems Division, Thiokol Corporation and MSFC personnel to study the feasibility of simulating the combustion species, temperatures and flow fields of a conventional solid rocket motor (SRM) with a versatile simulator system. The SSRCS design is based on hybrid rocket motor principles. The simulator uses a solid fuel and a gaseous oxidizer. Verification of the feasibility of a SSRCS system as a test bed was completed using flow field and system analyses, as well as empirical test data. A total of 27 hot firings of a subscale SSRCS motor were conducted at MSFC. Testing of the Small-scale SSRCS program was completed in October 1992. This paper, a compilation of reports from the above team members and additional analysis of the instrumentation results, will discuss the final results of the analyses and test programs.

  5. Development of Detonation Modeling Capabilities for Rocket Test Facilities: Hydrogen-Oxygen-Nitrogen Mixtures

    NASA Technical Reports Server (NTRS)

    Allgood, Daniel C.

    2016-01-01

    The objective of the presented work was to develop validated computational fluid dynamics (CFD) based methodologies for predicting propellant detonations and their associated blast environments. Applications of interest were scenarios relevant to rocket propulsion test and launch facilities. All model development was conducted within the framework of the Loci/CHEM CFD tool due to its reliability and robustness in predicting high-speed combusting flow-fields associated with rocket engines and plumes. During the course of the project, verification and validation studies were completed for hydrogen-fueled detonation phenomena such as shock-induced combustion, confined detonation waves, vapor cloud explosions, and deflagration-to-detonation transition (DDT) processes. The DDT validation cases included predicting flame acceleration mechanisms associated with turbulent flame-jets and flow-obstacles. Excellent comparison between test data and model predictions were observed. The proposed CFD methodology was then successfully applied to model a detonation event that occurred during liquid oxygen/gaseous hydrogen rocket diffuser testing at NASA Stennis Space Center.

  6. Marshall Team Recreates Goddard Rocket

    NASA Technical Reports Server (NTRS)

    2003-01-01

    In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

  7. Marshall Team Recreates Goddard Rocket

    NASA Technical Reports Server (NTRS)

    2003-01-01

    In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

  8. Performance Comparison of Axisymmetric and Three-dimensional Hydrogen Film Coolant Injection in a 110N Hydrogen/oxygen Rocket

    NASA Technical Reports Server (NTRS)

    Arrington, Lynn A.; Reed, Brian D.

    1992-01-01

    An experimental performance comparison of two geometrically different fuel film coolant injection sleeves was conducted on a 110 N gaseous hydrogen/oxygen rocket. One sleeve had slots milled axially down the walls and the other had a smooth surface to give axisymmetric flow. The comparison was made to investigate a conclusion in an earlier study that attributed a performance underprediction to a symplifying modeling assumption of axisymmetric fuel film flow. The smooth sleeve had higher overall performance at one film coolant percentage and approximately the same or slightly better at another. The study showed that the lack of modeling of three-dimensional effects was not the cause of the performance underprediction as speculated in earlier analytical studies.

  9. Influence of the technique for injection of flue gas and the configuration of the swirl burner throat on combustion of gaseous fuel and formation of nitrogen oxides in the flame

    NASA Astrophysics Data System (ADS)

    Dvoinishnikov, V. A.; Khokhlov, D. A.; Knyaz'kov, V. P.; Ershov, A. Yu.

    2017-05-01

    How the points at which the flue gas was injected into the swirl burner and the design of the burner outlet influence the formation and development of the flame in the submerged space, as well as the formation of nitrogen oxides in the combustion products, have been studied. The object under numerical investigation is the flame of the GMVI combined (oil/gas) burner swirl burner fitted with a convergent, biconical, cylindrical, or divergent throat at the burner outlet with individual supply of the air and injection of the gaseous fuel through tubing. The burners of two designs were investigated; they differ by the absence or presence of an inlet for individual injection of the flue gas. A technique for numerical simulation of the flame based on the CFD methods widely used in research of this kind underlies the study. Based on the summarized results of the numerical simulation of the processes that occur in jet flows, the specific features of the aerodynamic pattern of the flame have been established. It is shown that the flame can be conventionally divided into several sections over its length in all investigations. The lengths of each of the sections, as well as the form of the fields of axial velocity, temperatures, concentrations of the fuel, oxygen, and carbon and nitrogen oxides, are different and determined by the design features of the burner, the flow rates of the agent, and the compositions of the latter in the burner ducts as well as the configuration of the burner throat and the temperature of the environment. To what degree the burner throat configuration and the techniques for injection of the flue gas at different ambient temperatures influence the formation of nitrogen oxides has been established. It is shown that the supply of the recirculation of flue gas into the fuel injection zone enables a considerable reduction in the formation of nitrogen oxides in the flame combustion products. It has been established that the locations of the zones of

  10. Gaseous emissions from waste combustion.

    PubMed

    Werther, Joachim

    2007-06-18

    An overview is given on methods and technologies for limiting the gaseous emissions from waste combustion. With the guideline 2000/76/EC recent European legislation has set stringent limits not only for the mono-combustion of waste in specialized incineration plants but also for co-combustion in coal-fired power plants. With increased awareness of environmental issues and stepwise decrease of emission limits and inclusion of more and more substances into the network of regulations a multitude of emission abatement methods and technologies have been developed over the last decades. The result is the state-of-the-art waste incinerator with a number of specialized process steps for the individual components in the flue gas. The present work highlights some new developments which can be summarized under the common goal of reducing the costs of flue gas treatment by applying systems which combine the treatment of several noxious substances in one reactor or by taking new, simpler routes instead of the previously used complicated ones or - in the case of flue gas desulphurisation - by reducing the amount of limestone consumption. Cost reduction is also the driving force for new processes of conditioning of nonhomogenous waste before combustion. Pyrolysis or gasification is used for chemical conditioning whereas physical conditioning means comminution, classification and sorting processes. Conditioning yields a fuel which can be used in power plants either as a co-fuel or a mono-fuel and which will burn there under much better controlled conditions and therefore with less emissions than the nonhomogeneous waste in a conventional waste incinerator. Also for cost reasons, co-combustion of wastes in coal-fired power stations is strongly pressing into the market. Recent investigations reveal that the co-firing of waste can also have beneficial effects on the operating behavior of the boiler and on the gaseous emissions.

  11. CFD Analysis of the 24-inch JIRAD Hybrid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Liang, Pak-Yan; Ungewitter, Ronald; Claflin, Scott

    1996-01-01

    A series of multispecies, multiphase computational fluid dynamics (CFD) analyses of the 24-inch diameter joint government industry industrial research and development (JIRAD) hybrid rocket motor is described. The 24-inch JIRAD hybrid motor operates by injection of liquid oxygen (LOX) into a vaporization plenum chamber upstream of ports in the hydroxyl-terminated polybutadiene (HTPB) solid fuel. The injector spray pattern had a strong influence on combustion stability of the JIRAD motor so a CFD study was initiated to define the injector end flow field under different oxidizer spray patterns and operating conditions. By using CFD to gain a clear picture of the flow field and temperature distribution within the JIRAD motor, it is hoped that the fundamental mechanisms of hybrid combustion instability may be identified and then suppressed by simple alterations to the oxidizer injection parameters such as injection angle and velocity. The simulations in this study were carried out using the General Algorithm for Analysis of Combustion SYstems (GALACSY) multiphase combustion codes. GALACSY consists of a comprehensive set of droplet dynamic submodels (atomization, evaporation, etc.) and a computationally efficient hydrocarbon chemistry package built around a robust Navier-Stokes solver optimized for low Mach number flows. Lagrangian tracking of dispersed particles describes a closely coupled spray phase. The CFD cases described in this paper represent various levels of simplification of the problem. They include: (A) gaseous oxygen with combusting fuel vapor blowing off the walls at various oxidizer injection angles and velocities, (B) gaseous oxygen with combusting fuel vapor blowing off the walls, and (C) liquid oxygen with combusting fuel vapor blowing off the walls. The study used an axisymmetric model and the results indicate that the injector design significantly effects the flow field in the injector end of the motor. Markedly different recirculation patterns are

  12. Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion Measurements

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan; Zakany, James S.

    1996-01-01

    A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face

  13. Advanced bioreactor systems for gaseous substrates: Conversion of synthesis gas to liquid fuels and removal of SO{sub x} and NO{sub x} from coal combustion gases

    SciTech Connect

    Selvaraj, P.T.; Kaufman, E.N.

    1995-06-01

    The purpose of the proposed research program is the development and demonstration of a new generation of gaseous substrate-based bioreactors for the production of liquid fuels from coal synthesis gas and the removal of NO{sub x} and SO{sub x} species from combustion flue gas. Coal is thermochemically converted to synthesis gas consisting of carbon monoxide, hydrogen, and carbon dioxide. Conventional catalytic upgrading of coal synthesis gas into alcohols or other oxychemicals is subject to several processing problems such as interference of the other constituents in the synthesis gases, strict CO/H{sub 2} ratios required to maintain a particular product distribution and yield, and high processing cost due to the operation at high temperatures and pressures. Recently isolated and identified bacterial strains capable of utilizing CO as a carbon source and coverting CO and H{sub 2} into mixed alcohols offer the potential of performing synthesis gas conversion using biocatalysts. Biocatalytic conversion, though slower than the conventional process, has several advantages such as decreased interference of the other constituents in the synthesis gases, no requirement for strict CO/H{sub 2} ratios, and decreased capital and oeprating costs as the biocatalytic reactions occur at ambient temperatures and pressures.

  14. Hybrid rocket engine, theoretical model and experiment

    NASA Astrophysics Data System (ADS)

    Chelaru, Teodor-Viorel; Mingireanu, Florin

    2011-06-01

    The purpose of this paper is to build a theoretical model for the hybrid rocket engine/motor and to validate it using experimental results. The work approaches the main problems of the hybrid motor: the scalability, the stability/controllability of the operating parameters and the increasing of the solid fuel regression rate. At first, we focus on theoretical models for hybrid rocket motor and compare the results with already available experimental data from various research groups. A primary computation model is presented together with results from a numerical algorithm based on a computational model. We present theoretical predictions for several commercial hybrid rocket motors, having different scales and compare them with experimental measurements of those hybrid rocket motors. Next the paper focuses on tribrid rocket motor concept, which by supplementary liquid fuel injection can improve the thrust controllability. A complementary computation model is also presented to estimate regression rate increase of solid fuel doped with oxidizer. Finally, the stability of the hybrid rocket motor is investigated using Liapunov theory. Stability coefficients obtained are dependent on burning parameters while the stability and command matrixes are identified. The paper presents thoroughly the input data of the model, which ensures the reproducibility of the numerical results by independent researchers.

  15. Metallized Gelled Propellants: Oxygen/RP-1/aluminum Rocket Combustion Experiments

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan; Zakany, James S.

    1995-01-01

    A series of combustion experiments were conducted to measure the specific impulse, Cstar-, and specific-impulse efficiencies of a rocket engine using metallized gelled liquid propellants. These experiments used a small 20- to 40-1bf (89- to 178-N) thrust, modular engine consisting of an injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum and gaseous oxygen was the oxidizer. Ten different injectors were used during the testing: 6 for the baseline 02/RP-1 tests and 4 for the gelled fuel tests which covered a wide range of mixture ratios. At the peak of the Isp versus oxidizer-to-fuel ratio (O/F) data, a range of 93 to 99% Cstar efficiency was reached with ungelled 02/RP-1. A Cstar efficiency range of 75 to 99% was obtained with gelled RP-l (0-wt% RP-1/Al) while the metallized 5-wt% RP-1/Al delivered a Cstar efficiency of 94 to 99% at the peak Isp in the O/F range tested. An 88 to 99% Cstar efficiency was obtained at the peak Isp of the gelled RP1/Al with 55-wt% Al. Specific impulse efficiencies for the 55-wt% RP-1/Al of 67%-83% were obtained at a 2.4:1 expansion ratio. Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0- and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling. These experiments represent a first step in characterizing the performance of and operational issues with gelled RP-1 fuels.

  16. Atomic hydrogen rocket engine

    NASA Technical Reports Server (NTRS)

    Etters, R. D.; Flurchick, K.

    1981-01-01

    A rocket using atomic hydrogen propellant is discussed. An essential feature of the proposed engine is that the atomic hydrogen fuel is used as it is produced, thus eliminating the necessity of storage. The atomic hydrogen flows into a combustion chamber and recombines, producing high velocity molecular hydrogen which flows out an exhaust port. Standard thermodynamics, kinetic theory and wall recombination cross-sections are used to predict a thrust of approximately 1.4 N for a RF hydrogen flow rate of 4 x 10 to the 22nd/sec. Specific impulses are nominally from 1000 to 2000 sec. It is predicted that thrusts on the order of one Newton and specific impulses of up to 2200 sec are attainable with nominal RF discharge fluxes on the order of 10 to the 22nd atoms/sec; further refinements will probably not alter these predictions by more than a factor of two.

  17. Preliminary Experimental Investigation on Detonation Initiation in the Ejector of a Pulse Detonation Rocket Engine

    NASA Astrophysics Data System (ADS)

    Yan, Yu; Fan, Wei; Mu, Yang

    2012-12-01

    A small pulse detonation rocket engine (PDRE) was used as a predetonator to initiate detonation in its ejector. The detonation products discharged from the PDRE was not only ignition source for the ejector but also primary flow which entrained air from environment into the ejector. Stoichiometric liquid kerosene and gaseous oxygen were used as reactants for the PDRE. While in the ejector injected liquid kerosene was used as fuel and entrained air was used as oxidizer. Reactants in the ejector were ignited by the detonation wave and products discharged from the PDRE. Detonation was successfully initiation in present experiments. It was found that flame propagation upstream at the entrance of the ejector was inevitable, which affected the detonation initiation process in the ejector. Disks with orifices were placed at the entrance of the ejector to weaken the flame propagation upstream effect, which would affect the air flow entraining process, but the results show it worked.

  18. Air-breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This Quick Time movie depicts the Rocketdyne static test of an air-breathing rocket. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's advanced Transportation Program at the Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  19. Development of Advanced Hydrocarbon Fuels at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Bai, S. D.; Dumbacher, P.; Cole, J. W.

    2002-01-01

    This was a small-scale, hot-fire test series to make initial measurements of performance differences of five new liquid fuels relative to rocket propellant-1 (RP-1). The program was part of a high-energy-density materials development at Marshall Space Flight Center (MSFC), and the fuels tested were quadricyclane, 1-7 octodiyne, AFRL-1, biclopropylidene, and competitive impulse noncarcinogenic hypergol (CINCH) (di-methyl-aminoethyl-azide). All tests were conducted at MSFC. The first four fuels were provided by the U.S. Air Force Research Laboratory (AFRL), Edwards Air Force Base, CA. The U.S. Army, Redstone Arsenal, Huntsville, AL, provided the CINCH. The data recorded in all hot-fire tests were used to calculate specific impulse and characteristic exhaust velocity for each fuel, then compared to RP-1 at the same conditions. This was not an exhaustive study, comparing each fuel to RP-1 at an array of mixture ratios, nor did it include important fuel parameters, such as fuel handling or long-term storage. The test hardware was designed for liquid oxygen (lox)/RP-1, then modified for gaseous oxygen/RP-1 to avoid two-phase lox at very small flow rates. All fuels were tested using the same thruster/injector combination designed for RP-1. The results of this test will be used to determine which fuels will be tested in future test programs.

  20. The performance of a boron-loaded gel-fuel ramjet

    NASA Astrophysics Data System (ADS)

    Haddad, A.; Natan, B.; Arieli, R.

    2011-10-01

    The present work focuses on the possibility of combining the advantages of ramjet propulsion with the high energetic potential of boron. However, the use of boron poses two major challenges. The first, common to all solid additives to liquid fuels is particle sedimentation and poor dispersion. This problem is solved through the use of a gel fuel. The second obstacle, specific to boron-enriched fuels, is the difficulty in realizing the full energetic potential of boron. This could be overcome by means of an aft-combustion chamber, where fuel-rich combustion products are mixed with cold bypass air. Cooling causes the gaseous boron oxide to condense and, as a consequence, the heat of evaporation trapped in the gaseous oxide is released. The merits of such a combination are assessed through its ability to power an air-to-surface missile of relatively small size, capable of delivering a large payload to over a distance of about 1000 km in short time. The paper presents a preliminary design of a ramjet missile using a gel fuel loaded with boron. The thermochemical aspects of the two-stage combustion of the fuel are considered. A comparison with a solid rocket motor (SRM) missile launched under the same conditions as the ramjet missile is made. The boron-loaded gel-fuel ramjet is found superior for this mission.

  1. Radial flow nuclear thermal rocket (RFNTR)

    DOEpatents

    Leyse, Carl F.

    1995-11-07

    A radial flow nuclear thermal rocket fuel assembly includes a substantially conical fuel element having an inlet side and an outlet side. An annular channel is disposed in the element for receiving a nuclear propellant, and a second, conical, channel is disposed in the element for discharging the propellant. The first channel is located radially outward from the second channel, and separated from the second channel by an annular fuel bed volume. This fuel bed volume can include a packed bed of loose fuel beads confined by a cold porous inlet frit and a hot porous exit frit. The loose fuel beads include ZrC coated ZrC-UC beads. In this manner, nuclear propellant enters the fuel assembly axially into the first channel at the inlet side of the element, flows axially across the fuel bed volume, and is discharged from the assembly by flowing radially outward from the second channel at the outlet side of the element.

  2. Radial flow nuclear thermal rocket (RFNTR)

    DOEpatents

    Leyse, Carl F.

    1995-01-01

    A radial flow nuclear thermal rocket fuel assembly includes a substantially conical fuel element having an inlet side and an outlet side. An annular channel is disposed in the element for receiving a nuclear propellant, and a second, conical, channel is disposed in the element for discharging the propellant. The first channel is located radially outward from the second channel, and separated from the second channel by an annular fuel bed volume. This fuel bed volume can include a packed bed of loose fuel beads confined by a cold porous inlet frit and a hot porous exit frit. The loose fuel beads include ZrC coated ZrC-UC beads. In this manner, nuclear propellant enters the fuel assembly axially into the first channel at the inlet side of the element, flows axially across the fuel bed volume, and is discharged from the assembly by flowing radially outward from the second channel at the outlet side of the element.

  3. Fluidized-Solid-Fuel Injection Process

    NASA Technical Reports Server (NTRS)

    Taylor, William

    1992-01-01

    Report proposes development of rocket engines burning small grains of solid fuel entrained in gas streams. Main technical discussion in report divided into three parts: established fluidization technology; variety of rockets and rocket engines used by nations around the world; and rocket-engine equation. Discusses significance of specific impulse and ratio between initial and final masses of rocket. Concludes by stating three important reasons to proceed with new development: proposed engines safer; fluidized-solid-fuel injection process increases variety of solid-fuel formulations used; and development of fluidized-solid-fuel injection process provides base of engineering knowledge.

  4. 78. PIPING CHANNEL FOR FUEL LOADING, FUEL TOPPING, COMPRESSED AIR, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    78. PIPING CHANNEL FOR FUEL LOADING, FUEL TOPPING, COMPRESSED AIR, GASEOUS NITROGEN, AND HELIUM - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 East, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

  5. Metal hydride and pyrophoric fuel additives for dicyclopentadiene based hybrid propellants

    NASA Astrophysics Data System (ADS)

    Shark, Steven C.

    between 80% and 90%. The regression rate and C* efficiency mass flux dependence indicate a shift towards a more diffusion controlled system with metal hydride particle addition. Although these types of energetic particles have potential as high performing fuel additives, they can be in low supply and expensive. An opposed flow burner was investigated as a means to screen and characterize hybrid rocket fuels prior to full scale rocket motor testing. Although this type of configuration has been investigated in the past, no comparison has been made to hybrid rocket motor operation in terms of mass flux. Polymeric fuels and low melt temperature fuels with and without additives were investigated via an opposed flow burner. The effects of laminar and turbulent flow regimes on the convective heat transfer in the opposed flow system was depicted in the regression rate trends of these fuels. Regression rate trends similar to hybrid rocket motor operation were depicted, including the entrainment mechanism for paran fuel. However, there was a shift in overall magnitude of these results. A decrease in regression rate occurred for HTPB loaded with passivated nano-aluminum, due to low resonance time in the reaction zone. Previous results have shown that pyrophoric additives can cause an increase in regression rate in the opposed flow burner configuration. It is proposed that the opposed burner is useful as a screening and characterization tool for some propellant combinations. Gaseous oxygen (GOX) was investigated as an oxidizer for similar fuels evaluated with RGHP. Specifically, combustion performance sensitivity to mass flux and MH particle size was investigated. Similar results to the RGHP experiments were observed for the regression rate tends of HTPB, DPCD, and NabH 4 addition. Kinetically limited regression rate dependence on mass flux was observed at the higher mass flux levels. No major increase in C* efficiency was observed for MH addition. The C* efficiency varied with

  6. Solar Thermal Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Sercel, J. C.

    1986-01-01

    Paper analyzes potential of solar thermal rockets as means of propulsion for planetary spacecraft. Solar thermal rocket uses concentrated Sunlight to heat working fluid expelled through nozzle to produce thrust.

  7. American Rocket Society

    NASA Technical Reports Server (NTRS)

    2004-01-01

    In addition to Dr. Robert Goddard's pioneering work, American experimentation in rocketry prior to World War II grew, primarily in technical societies. This is an early rocket motor designed and developed by the American Rocket Society in 1932.

  8. ROBOTS TO ROCKET CITY

    NASA Image and Video Library

    2016-03-06

    HIGH SCHOOL STUDENTS FROM NORTH ALABAMA GATHER AT THE U.S. SPACE AND ROCKET CENTER'S DAVIDSON CENTER FOR THE "ROBOTS TO ROCKET CITY" EVENT SHOWCASING THEIR INDIVIDUAL ROBOTS PRIOR TO LATER COMPETITIONS.

  9. 148. SKID 2 FOR LOADING ROCKET PROPELLANT AT EAST SIDE ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    148. SKID 2 FOR LOADING ROCKET PROPELLANT AT EAST SIDE OF FUEL CONTROL ROOM (215), LSB (BLDG. 751) - Vandenberg Air Force Base, Space Launch Complex 3, Launch Pad 3 East, Napa & Alden Roads, Lompoc, Santa Barbara County, CA

  10. Gaseous wire detectors

    SciTech Connect

    Va'vra, J.

    1997-08-01

    This article represents a series of three lectures describing topics needed to understand the design of typical gaseous wire detectors used in large high energy physics experiments; including the electrostatic design, drift of electrons in the electric and magnetic field, the avalanche, signal creation, limits on the position accuracy as well as some problems one encounters in practical operations.

  11. Metallic Hydrogen: A Game Changing Rocket Propellant

    NASA Technical Reports Server (NTRS)

    Silvera, Isaac F.

    2016-01-01

    The objective of this research is to produce metallic hydrogen in the laboratory using an innovative approach, and to study its metastability properties. Current theoretical and experimental considerations expect that extremely high pressures of order 4-6 megabar are required to transform molecular hydrogen to the metallic phase. When metallic hydrogen is produced in the laboratory it will be extremely important to determine if it is metastable at modest temperatures, i.e. remains metallic when the pressure is released. Then it could be used as the most powerful chemical rocket fuel that exists and revolutionize rocketry, allowing single-stage rockets to enter orbit and chemically fueled rockets to explore our solar system.

  12. Transpiration cooled throat for hydrocarbon rocket engines

    NASA Technical Reports Server (NTRS)

    May, Lee R.; Burkhardt, Wendel M.

    1991-01-01

    The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.

  13. Heat transfer to throat tubes in a square-chambered rocket engine at the NASA Lewis Research Center

    NASA Technical Reports Server (NTRS)

    Nesbitt, James A.; Brindley, William J.

    1989-01-01

    A gaseous H2/O2 rocket engine was constructed at the NASA-Lewis to provide a high heat flux source representative of the heat flux to the blades in the high pressure fuel turbopump (HPFTP) during startup of the space shuttle main engines. The high heat flux source was required to evaluate the durability of thermal barrier coatings being investigated for use on these blades. The heat transfer, and specifically, the heat flux to tubes located at the throat of the test rocket engine was evaluated and compared to the heat flux to the blades in the HPFTP during engine startup. Gas temperatures, pressures and heat transfer coefficients in the test rocket engine were measured. Near surface metal temperatures below thin thermal barrier coatings were also measured at various angular orientations around the throat tube to indicate the angular dependence of the heat transfer coefficients. A finite difference model for a throat tube was developed and a thermal analysis was performed using the measured gas temperatures and the derived heat transfer coefficients to predict metal temperatures in the tube. Near surface metal temperatures of an uncoated throat tube were measured at the stagnation point and showed good agreement with temperatures predicted by the thermal model. The maximum heat flux to the throat tube was calculated and compared to that predicted for the leading edge of an HPFTP blade. It is shown that the heat flux to an uncooled throat tube is slightly greater than the heat flux to an HPFTP blade during engine startup.

  14. Effect of Soot Particles on Supersonic Rocket Plume Properties

    NASA Astrophysics Data System (ADS)

    Gaissinski, Igor; Levy, Yeshayahou; Lev, Mikhael; Sherbaum, Valery

    2012-06-01

    Plumes from hydrocarbon-fueled rockets usually contain some amount of soot. In spite of the small amount, such soot particles can play a critical role in the characteristics of the infrared radiation emission since soot radiates a continuous, near-blackbody spectrum. The contribution of the soot to the plume radiation depends on the amount of soot, the physical properties of the particles, their concentration, and their temperature distribution in the flow field. The trajectories of solid particles and their temperatures can differ from those of the gas due to the particle mechanical and thermal inertia. CFD FLUENT code for solving two-phase Navier-Stokes equations coupled with chemical reactions and soot particle combustion was applied for exhaust plume simulations. Exhaust plumes with soot mass loading of 2% were simulated for three altitudes of 2 km, 8 km and 16 km. Radial distributions of the cloud particle density were obtained for different distances downstream the exhaust nozzle. As a result of the particle deceleration at the boundary layer inside the nozzle the particle concentration increased at the plume periphery. The particle temperature was higher than the gaseous temperature of the plume. The temperature difference between the soot particle and gas along corresponding trajectories was about 5-10%. The infrared radiation from the plumes with carbon soot was calculated. Its intensity was found to be dependent on the particle distribution in the plume.

  15. The OGRESS Sounding Rocket Payload

    NASA Astrophysics Data System (ADS)

    Rogers, Thomas; Zeiger, B.; McEntaffer, R. L.; Schultz, T.; Oakley, P.; Cash, W. C.

    2013-01-01

    We present an overview of the Off-plane Grating Rocket for Extended Source Spectroscopy (OGRESS) sounding rocket payload based at the University of Iowa and University of Colorado, Boulder. The payload has been launched three times before under the names CyXESS, EXOS, and CODEX and is scheduled to fly again in February 2014. The payload is designed to observe large diffuse soft X-ray sources between ~100 - 1000 eV. OGRESS will observe the Vela supernova remnant and achieve the highest spectral resolution ever taken of this object in our bandpass. OGRESS does not use the standard optical design of a grazing incidence Wolter telescope. Instead, OGRESS uses a wire-grid collimator to remove any nonconverging light. This optical design leads to poor sensitivity to point sources, but works very well for large extended sources, for which the telescope beam is fully illuminated. The light is dispersed by an array of off-plane parallel-groove sinusoidal gratings and focused onto Gaseous Electron Multiplier (GEM) detectors.

  16. Fiber optics in liquid propellant rocket engine environments

    NASA Technical Reports Server (NTRS)

    Delcher, R.; Dinnsen, D.; Barkhoudarian, S.

    1991-01-01

    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.

  17. Determination of 1-methyl-1H-1,2,4-triazole in soils contaminated by rocket fuel using solid-phase microextraction, isotope dilution and gas chromatography-mass spectrometry.

    PubMed

    Yegemova, Saltanat; Bakaikina, Nadezhda V; Kenessov, Bulat; Koziel, Jacek A; Nauryzbayev, Mikhail

    2015-10-01

    Environmental monitoring of Central Kazakhstan territories where heavy space booster rockets land requires fast, efficient, and inexpensive analytical methods. The goal of this study was to develop a method for quantitation of the most stable transformation product of rocket fuel, i.e., highly toxic unsymmetrical dimethylhydrazine - 1-methyl-1H-1,2,4-triazole (MTA) in soils using solid-phase microextraction (SPME) in combination with gas chromatography-mass spectrometry. Quantitation of organic compounds in soil samples by SPME is complicated by a matrix effect. Thus, an isotope dilution method was chosen using deuterated analyte (1-(trideuteromethyl)-1H-1,2,4-triazole; MTA-d3) for matrix effect control. The work included study of the matrix effect, optimization of a sample equilibration stage (time and temperature) after spiking MTA-d3 and validation of the developed method. Soils of different type and water content showed an order of magnitude difference in SPME effectiveness of the analyte. Isotope dilution minimized matrix effects. However, proper equilibration of MTA-d3 in soil was required. Complete MTA-d3 equilibration at temperatures below 40°C was not observed. Increase of temperature to 60°C and 80°C enhanced equilibration reaching theoretical MTA/MTA-d3 response ratios after 13 and 3h, respectively. Recoveries of MTA depended on concentrations of spiked MTA-d3 during method validation. Lowest spiked MTA-d3 concentration (0.24 mg kg(-1)) provided best MTA recoveries (91-121%). Addition of excess water to soil sample prior to SPME increased equilibration rate, but it also decreased method sensitivity. Method detection limit depended on soil type, water content, and was always below 1 mg kg(-1). The newly developed method is fully automated, and requires much lower time, labor and financial resources compared to known methods. Copyright © 2015 Elsevier B.V. All rights reserved.

  18. How High? How Fast? How Long? Modeling Water Rocket Flight with Calculus

    ERIC Educational Resources Information Center

    Ashline, George; Ellis-Monaghan, Joanna

    2006-01-01

    We describe an easy and fun project using water rockets to demonstrate applications of single variable calculus concepts. We provide procedures and a supplies list for launching and videotaping a water rocket flight to provide the experimental data. Because of factors such as fuel expulsion and wind effects, the water rocket does not follow the…

  19. How High? How Fast? How Long? Modeling Water Rocket Flight with Calculus

    ERIC Educational Resources Information Center

    Ashline, George; Ellis-Monaghan, Joanna

    2006-01-01

    We describe an easy and fun project using water rockets to demonstrate applications of single variable calculus concepts. We provide procedures and a supplies list for launching and videotaping a water rocket flight to provide the experimental data. Because of factors such as fuel expulsion and wind effects, the water rocket does not follow the…

  20. Moving Material into Space Without Rockets.

    ERIC Educational Resources Information Center

    Cheng, R. S.; Trefil, J. S.

    1985-01-01

    In response to conventional rocket demands on fuel supplies, electromagnetic launches were developed to give payloads high velocity using a stationary energy source. Several orbital mechanics problems are solved including a simple problem (radial launch with no rotation) and a complex problem involving air resistance and gravity. (DH)

  1. Program For Optimization Of Nuclear Rocket Engines

    NASA Technical Reports Server (NTRS)

    Plebuch, R. K.; Mcdougall, J. K.; Ridolphi, F.; Walton, James T.

    1994-01-01

    NOP is versatile digital-computer program devoloped for parametric analysis of beryllium-reflected, graphite-moderated nuclear rocket engines. Facilitates analysis of performance of engine with respect to such considerations as specific impulse, engine power, type of engine cycle, and engine-design constraints arising from complications of fuel loading and internal gradients of temperature. Predicts minimum weight for specified performance.

  2. Moving Material into Space Without Rockets.

    ERIC Educational Resources Information Center

    Cheng, R. S.; Trefil, J. S.

    1985-01-01

    In response to conventional rocket demands on fuel supplies, electromagnetic launches were developed to give payloads high velocity using a stationary energy source. Several orbital mechanics problems are solved including a simple problem (radial launch with no rotation) and a complex problem involving air resistance and gravity. (DH)

  3. Rockets for spin recovery

    NASA Technical Reports Server (NTRS)

    Whipple, R. D.

    1980-01-01

    The potential effectiveness of rockets as an auxiliary means for an aircraft to effect recovery from spins was investigated. The advances in rocket technology produced by the space effort suggested that currently available systems might obviate many of the problems encountered in earlier rocket systems. A modern fighter configuration known to exhibit a flat spin mode was selected. An analytical study was made of the thrust requirements for a rocket spin recovery system for the subject configuration. These results were then applied to a preliminary systems study of rocket components appropriate to the problem. Subsequent spin tunnel tests were run to evaluate the analytical results.

  4. Rocket nozzle snubber

    NASA Astrophysics Data System (ADS)

    Hoever, Jeffrey P.; Reginato, Richard L.; Bay, Frederick M.; Rolon, Claudia E.; Bazigos, Michael

    1991-10-01

    Stress from the ignition of a rocket is transferred from the rocket nozzle to the actuators and flexible joint seal which move the nozzle causing potential failure of the actuators. This stress can be relieved by using a snubber in the form of a flexible hollow conical toroidal bladder filled with incompressible fluid fitted snugly between the nozzle and the main body of the rocket. Such a snubber transfers the stress to the main body of the rocket through the snubber while deforming to allow the rocket nozzle to slew in response to control commands.

  5. Vented nuclear fuel element

    DOEpatents

    Grossman, Leonard N.; Kaznoff, Alexis I.

    1979-01-01

    A nuclear fuel cell for use in a thermionic nuclear reactor in which a small conduit extends from the outside surface of the emitter to the center of the fuel mass of the emitter body to permit escape of volatile and gaseous fission products collected in the center thereof by virtue of molecular migration of the gases to the hotter region of the fuel.

  6. Uranium droplet core nuclear rocket

    NASA Technical Reports Server (NTRS)

    Anghaie, Samim

    1991-01-01

    Uranium droplet nuclear rocket is conceptually designed to utilize the broad temperature range ofthe liquid phase of metallic uranium in droplet configuration which maximizes the energy transfer area per unit fuel volume. In a baseline system dissociated hydrogen at 100 bar is heated to 6000 K, providing 2000 second of Isp. Fission fragments and intense radian field enhance the dissociation of molecular hydrogen beyond the equilibrium thermodynamic level. Uranium droplets in the core are confined and separated by an axisymmetric vortex flow generated by high velocity tangential injection of hydrogen in the mid-core regions. Droplet uranium flow to the core is controlled and adjusted by a twin flow nozzle injection system.

  7. Uranium droplet core nuclear rocket

    NASA Technical Reports Server (NTRS)

    Anghaie, Samim

    1991-01-01

    Uranium droplet nuclear rocket is conceptually designed to utilize the broad temperature range ofthe liquid phase of metallic uranium in droplet configuration which maximizes the energy transfer area per unit fuel volume. In a baseline system dissociated hydrogen at 100 bar is heated to 6000 K, providing 2000 second of Isp. Fission fragments and intense radian field enhance the dissociation of molecular hydrogen beyond the equilibrium thermodynamic level. Uranium droplets in the core are confined and separated by an axisymmetric vortex flow generated by high velocity tangential injection of hydrogen in the mid-core regions. Droplet uranium flow to the core is controlled and adjusted by a twin flow nozzle injection system.

  8. Lagrangian modeling of turbulent spray combustion: application to rocket engines cryogenic conditions

    NASA Astrophysics Data System (ADS)

    Izard, J.-F.; Mura, A.

    2011-10-01

    The present work is concerned with the application of a turbulent two-phase flow combustion model to a spray flame of Liquid Oxygen (LOx) and Gaseous Hydrogen (GH2). The proposed strategy relies on a joint Eulerian-Lagrangian framework. The Probability Density Function (PDF) that characterizes the liquid phase is evaluated by simulating the Williams spray equation [1] thanks to the semifluid approach introduced in [2]. The Lagrangian approach provides the classical exchange terms with the gaseous phase and, especially, several vaporization source terms. They are required to describe turbulent combustion but difficult to evaluate from the Eulerian point of view. The turbulent combustion model retained here relies on the consideration of the mixture fraction to evaluate the local fuel-to-oxidizer ratio, and the oxygen mass fraction to follow the deviations from chemical equilibrium. The difficulty associated with the estimation of a joint scalar PDF is circumvented by invoking the sudden chemistry hypothesis [3]. In this manner, the problem reduces to the estimation of the mixture fraction PDF, but with the influence of the terms related to vaporization that are the source of additional fluctuations of composition. Following the early proposal of [4], these terms are easily obtained from the Lagrangian framework adopted to describe the two-phase flows. The resulting computational model is applied to the numerical simulation of LOx-GH2 spray flames. The test case (Mascotte) is representative of combustion in rocket engine conditions. The results of numerical simulations display a satisfactory agreement with available experimental data.

  9. Theoretical and experimental analysis of liquid layer instability in hybrid rocket engines

    NASA Astrophysics Data System (ADS)

    Kobald, Mario; Verri, Isabella; Schlechtriem, Stefan

    2015-03-01

    The combustion behavior of different hybrid rocket fuels has been analyzed in the frame of this research. Tests have been performed in a 2D slab burner configuration with windows on two sides. Four different liquefying paraffin-based fuels, hydroxyl terminated polybutadiene (HTPB) and high-density polyethylene (HDPE) have been tested in combination with gaseous oxygen (GOX). Experimental high-speed video data have been analyzed manually and with the proper orthogonal decomposition (POD) technique. Application of POD enables the recognition of the main structures of the flow field and the combustion flame appearing in the video data. These results include spatial and temporal analysis of the structures. For liquefying fuels these spatial values relate to the wavelengths associated to the Kelvin Helmholtz Instability (KHI). A theoretical long-wave solution of the KHI problem shows good agreement with the experimental results. Distinct frequencies found in the POD analysis can be related to the precombustion chamber configuration which can lead to vortex shedding phenomena.

  10. GASEOUS DISPOSAL PROCESS

    DOEpatents

    Ryan, R.F.; Thomasson, F.R.; Hicks, J.H.

    1963-01-22

    A method is described of removing gaseous radioactive Xe and Kr from water containing O. The method consists in stripping the gases from the water stream by means of H flowing countercurrently to the stream. The gases are then heated in a deoxo bed to remove O. The carrier gas is next cooled and passed over a charcoal adsorbent bed maintained at a temperature of about --280 deg F to remove the Xe and Kr. (AEC)

  11. Gaseous diffusion system

    DOEpatents

    Garrett, George A.; Shacter, John

    1978-01-01

    1. A gaseous diffusion system comprising a plurality of diffusers connected in cascade to form a series of stages, each of said diffusers having a porous partition dividing it into a high pressure chamber and a low pressure chamber, and means for combining a portion of the enriched gas from a succeeding stage with a portion of the enriched gas from the low pressure chamber of each stage and feeding it into one extremity of the high pressure chamber thereof.

  12. Regression rate study of porous axial-injection, endburning hybrid fuel grains

    NASA Astrophysics Data System (ADS)

    Hitt, Matthew A.

    This experimental and theoretical work examines the effects of gaseous oxidizer flow rates and pressure on the regression rates of porous fuels for hybrid rocket applications. Testing was conducted using polyethylene as the porous fuel and both gaseous oxygen and nitrous oxide as the oxidizer. Nominal test articles were tested using 200, 100, 50, and 15 micron fuel pore sizes. Pressures tested ranged from atmospheric to 1160 kPa for the gaseous oxygen tests and from 207 kPa to 1054 kPa for the nitrous oxide tests, and oxidizer injection velocities ranged from 35 m/s to 80 m/s for the gaseous oxygen tests and from 7.5 m/s to 16.8 m/s for the nitrous oxide tests. Regression rates were determined using pretest and posttest length measurements of the solid fuel. Experimental results demonstrated that the regression rate of the porous axial-injection, end-burning hybrid was a function of the chamber pressure, as opposed to the oxidizer mass flux typical in conventional hybrids. Regression rates ranged from approximately 0.75 mm/s at atmospheric pressure to 8.89 mm/s at 1160 kPa for the gaseous oxygen tests and 0.21 mm/s at 207 kPa to 1.44 mm/s at 1054 kPa for the nitrous oxide tests. The analytical model was developed based on a standard ablative model modified to include oxidizer flow through the grain. The heat transfer from the flame was primarily modeled using an empirically determined flame coefficient that included all heat transfer mechanisms in one term. An exploratory flame model based on the Granular Diffusion Flame model used for solid rocket motors was also adapted for comparison with the empirical flame coefficient. This model was then evaluated quantitatively using the experimental results of the gaseous oxygen tests as well as qualitatively using the experimental results of the nitrous oxide tests. The model showed agreement with the experimental results indicating it has potential for giving insight into the flame structure in this motor configuration

  13. Scaling a single element combustor to replicate combustion instability modes of a liquid rocket engine

    NASA Astrophysics Data System (ADS)

    Sweeney, Brian A.

    This research evaluated a method of scaling a single element sub-scale combustor to match the combustion instability modes of a full-scale liquid rocket engine. The experiments used a shear-coaxial injector in an atmospheric chamber using gaseous oxygen and a heated gaseous methane/nitrogen fuel mixture. The flow conditions matched the full-scale equivalence ratio, propellant velocities and propellant volumetric flow rates. The first set of experiments empirically determined the effect of chamber diameter on chamber temperature. The results were used to calculate the dimensions of the sub-scaled combustion chamber that would match the transverse frequencies of the full-scale engine. The scaled chamber was used in two sets of experiments. The stationary tests placed the injector at the center of the chamber and 0.25 in. from the wall. The centered test displayed evidence of coupling between the 1L chamber mode and the injector oxygen post at 885 Hz. Injector coupling was also observed during experiments with the full-scale rocket engine. With the injector 0.25 in. from the wall, the average chamber temperature dropped about 350°C from the centered test. As a consequence, the frequencies of the transverse modes were lower than the full-scale values. No major difference was found in this research between the stable and unstable set points of the full-scale engine. A translating stage was used to evaluate where various chamber modes appear as a function of injector location. The results show that the 1L chamber mode is present at every location and transverse modes appear as the injector moves near the wall.

  14. Researcher Poses with a Nuclear Rocket Model

    NASA Image and Video Library

    1961-11-21

    A researcher at the NASA Lewis Research Center with slide ruler poses with models of the earth and a nuclear-propelled rocket. The Nuclear Engine for Rocket Vehicle Applications (NERVA) was a joint NASA and Atomic Energy Commission (AEC) endeavor to develop a nuclear-powered rocket for both long-range missions to Mars and as a possible upper-stage for the Apollo Program. The early portion of the program consisted of basic reactor and fuel system research. This was followed by a series of Kiwi reactors built to test nuclear rocket principles in a non-flying nuclear engine. The next phase, NERVA, would create an entire flyable engine. The AEC was responsible for designing the nuclear reactor and overall engine. NASA Lewis was responsible for developing the liquid-hydrogen fuel system. The nuclear rocket model in this photograph includes a reactor at the far right with a hydrogen propellant tank and large radiator below. The payload or crew would be at the far left, distanced from the reactor.

  15. Scaled Rocket Testing in Hypersonic Flow

    NASA Technical Reports Server (NTRS)

    Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish

    2015-01-01

    NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

  16. Low thrust chemical rocket technology

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    1992-01-01

    An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher

  17. Theoretical performance of JP-4 fuel with a 70-30 mixture of fluorine and oxygen as a rocket propellant : equilibrium composition

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Huff, Vearl N

    1956-01-01

    Data were calculated for equivalence ratios of 1 to 4, chamber pressures of 300 and 600 pounds per square inch absolute, and pressure ratios of 1 to 1500. Parameters included are specific impulse, combustion and exit temperatures, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. A correlation is given which permits determination of performance for a wide range of chamber pressures. A method for obtaining specific impulse of JP-4 fuel with OF2 and O3-F2 mixtures is given.

  18. Nuclear thermal rockets using indigenous Martian propellants

    SciTech Connect

    Zubrin, R.M.

    1989-01-01

    This paper considers a novel concept for a Martian descent and ascent vehicle, called NIMF (for nuclear rocket using indigenous Martian fuel), the propulsion for which will be provided by a nuclear thermal reactor which will heat an indigenous Martian propellant gas to form a high-thrust rocket exhaust. The performance of each of the candidate Martian propellants, which include CO2, H2O, CH4, N2, CO, and Ar, is assessed, and the methods of propellant acquisition are examined. Attention is also given to the issues of chemical compatibility between candidate propellants and reactor fuel and cladding materials, and the potential of winged Mars supersonic aircraft driven by this type of engine. It is shown that, by utilizing the nuclear landing craft in combination with a hydrogen-fueled nuclear thermal interplanetary vehicle and a heavy lift booster, it is possible to achieve a manned Mars mission in one launch. 6 refs.

  19. A QSAR/QSTR study on the human health impact of the rocket fuel 1,1-dimethyl hydrazine and its transformation products Multicriteria hazard ranking based on partial order methodologies.

    PubMed

    Carlsen, Lars; Kenessov, Bulat N; Batyrbekova, Svetlana Ye

    2009-05-01

    The possible impact of the rocket fuel 1,1-dimethyl hydrazine (heptyl) (1) and its transformation products on human health has been studied using (Quantitative) Structure Activity/Toxicity ((Q)SAR/(Q)STR) modelling, including both ADME models and models for acute toxicity, organ specific adverse haematological effects, the cardiovascular and gastrointestinal systems, the kidneys, the liver and the lungs, as well as a model predicting the biological activity of the compounds. It was predicted that all compounds studied are readily bioavailable through oral intake and that significant amounts of the compounds will be freely available in the systemic circulation. In general, the compounds are not predicted to be acutely toxic apart from hydrogen cyanide, whereas several compounds are predicted to cause adverse organ specific human health effects. Further, several compounds are predicted to exhibit high probabilities for potential carcinogenicity, mutagenicity, teratogenicity and/or embryotoxicity. The compounds were ranked based on their predicted human health impact using partial order ranking methodologies that highlight which compounds on a cumulative basis should receive the major attention, i.e., N-nitroso dimethyl amine, 1,1,4,4-tetramethyl tetrazene, trimethyl, trimethyl hydrazine, acetaldehyde dimethyl hydrazone, 1, 1-formyl 2,2-dimethyl hydrazine and formaldehyde dimethyl hydrazone, respectively. Copyright © 2009 Elsevier B.V. All rights reserved.

  20. Highlights of 50 years of Aerojet, a pioneering American rocket company, 1942-1992

    NASA Astrophysics Data System (ADS)

    Winter, Frank H.; James, George S.

    1995-05-01

    The "pre-history" of Aerojet is recalled, followed by a survey of Aerojet's solid-fuel and liquid-fuel JATOs (Jet-Assisted Take-Off) to aircraft prime powerplants, missile sustainer motors, boosters, sounding rocket engines and, finally, nuclear powered rocket engines (NERVA).