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Sample records for gaseous rocket fuels

  1. Experimental study of rocket engine model with gaseous polyethylene fuel

    NASA Astrophysics Data System (ADS)

    Yemets, V. V.

    Experimental results for liquid rocket engine models with gaseous polyethylene fuel that is hard before its consumption are considered. The possibility of hard design element combustion in a liquid rocket engine is demonstrated.

  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. 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.

  5. 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.

  6. Gaseous fuel reactor systems for aerospace applications

    NASA Technical Reports Server (NTRS)

    Thom, K.; Schwenk, F. C.

    1977-01-01

    Research on the gaseous fuel nuclear rocket concept continues under the programs of the U.S. National Aeronautics and Space Administration (NASA) Office for Aeronautics and Space Technology and now includes work related to power applications in space and on earth. In a cavity reactor test series, initial experiments confirmed the low critical mass determined from reactor physics calculations. Recent work with flowing UF6 fuel indicates stable operation at increased power levels. Preliminary design and experimental verification of test hardware for high-temperature experiments have been accomplished. Research on energy extraction from fissioning gases has resulted in lasers energized by fission fragments. Combined experimental results and studies indicate that gaseous-fuel reactor systems have significant potential for providing nuclear fission power in space and on earth.

  7. 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.

  8. Measurements of reactive gaseous rocket injector response factors

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    The results presented represent the first successful attempt at the measurement of the driving capabilities of coaxial gaseous propellant rocket injectors. The required data have been obtained by employing the modified impedance tube technique with compressed air as the oxidizer and acetylene gas as the fuel. The data describe the frequency dependence of the injector admittances, from which the frequency dependence of the injection response factors can be calculated. The measured injector admittances have been compared with the predictions of the Feiler and Heidmann (1967) analytical model assuming different values for the characteristic combustion time. The values of combustion time which result in a best fit between the measured and predicted data are indicated for different equivalence ratios. It is shown that for the coaxial injector system investigated in this study the characteristic combustion times vary between .7 and 1.2 msec for equivalence ratios in the range of .57 to 1.31. The experimental data clearly show that the tested injector system could indeed drive combustion instabilities over a frequency range that is in qualitative agreement with the predictions of the Feiler and Heidmann model.

  9. 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

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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

  15. Study Of Fuels For Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Strand, Leon D.; Ray, Robert L.; Anderson, Floyd A.; Cohen, Norman S.

    1994-01-01

    Report describes experimental study of combustion and rates of regression of selected fuels for hybrid rocket engines. Part of continuing effort to develop fuels with greater rates of regression and lesser dependence on shapes of fuel grains and to maximize potential specific impulse at low cost.

  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. 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...

  18. 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...

  19. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL... Measurements for Gaseous Fuels per 100 Standard Cubic Feet Fuel Gallon equivalent measurement...

  20. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL... Measurements for Gaseous Fuels per 100 Standard Cubic Feet Fuel Gallon equivalent measurement...

  1. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL... Measurements for Gaseous Fuels per 100 Standard Cubic Feet Fuel Gallon equivalent measurement...

  2. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL... Measurements for Gaseous Fuels per 100 Standard Cubic Feet Fuel Gallon equivalent measurement...

  3. 49 CFR 538.8 - Gallon Equivalents for Gaseous Fuels.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... TRAFFIC SAFETY ADMINISTRATION, DEPARTMENT OF TRANSPORTATION MANUFACTURING INCENTIVES FOR ALTERNATIVE FUEL... Measurements for Gaseous Fuels per 100 Standard Cubic Feet Fuel Gallon equivalent measurement...

  4. Solid-Fuel Regression Rate for Standard-Flow Hybrid Rocket Motors

    NASA Astrophysics Data System (ADS)

    Morita, Takakazu; Yuasa, Saburo; Yamaguchi, Shigeru; Shimada, Toru

    Marxman's diffusion-limited analysis of hybrid rocket combustion has been often used to investigate various combustion problems in hybrid rocket motors. This analysis was developed on the basis of the Reynolds analogy in turbulent boundary layers. This analogy assumes that both molecular and turbulent Prandtl numbers are equal to one. In the present study, a semi-empirical correlation between the Stanton number and the skin-friction coefficient in a turbulent boundary layer was obtained. This is applicable to hybrid rocket combustion, and also includes the effects of the Prandtl numbers variation. Using this correlation, a fuel regression rate equation for standard-flow hybrid rocket motors was obtained, and its characteristics were examined. In addition, the calculated regression rate characteristics were compared with the experimental data from the laboratory-scale hybrid rocket motors that used gaseous oxygen (GOX) as oxidizer and polymethylmethacrylate (PMMA) as fuel.

  5. 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.

  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. 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.

  8. Gaseous fuel reactors for power systems

    NASA Technical Reports Server (NTRS)

    Helmick, H. H.; Schwenk, F. C.

    1978-01-01

    The Los Alamos Scientific Laboratory is participating in a NASA-sponsored program to demonstrate the feasibility of a gaseous uranium fueled reactor. The work is aimed at acquiring experimental and theoretical information for the design of a prototype plasma core reactor which will test heat removal by optical radiation. The basic goal of this work is for space applications, however, other NASA-sponsored work suggests several attractive applications to help meet earth-bound energy needs. Such potential benefits are: small critical mass, on-site fuel processing, high fuel burnup, low fission fragment inventory in reactor core, high temperature for process heat, optical radiation for photochemistry and space power transmission, and high temperature for advanced propulsion systems.

  9. 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.

  10. 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.

  11. Prediction of pressure and flow transients in a gaseous bipropellant reaction control rocket engine

    NASA Technical Reports Server (NTRS)

    Markowsky, J. J.; Mcmanus, H. N., Jr.

    1974-01-01

    An analytic model is developed to predict pressure and flow transients in a gaseous hydrogen-oxygen reaction control rocket engine feed system. The one-dimensional equations of momentum and continuity are reduced by the method of characteristics from partial derivatives to a set of total derivatives which describe the state properties along the feedline. System components, e.g., valves, manifolds, and injectors are represented by pseudo steady-state relations at discrete junctions in the system. Solutions were effected by a FORTRAN IV program on an IBM 360/65. The results indicate the relative effect of manifold volume, combustion lag time, feedline pressure fluctuations, propellant temperature, and feedline length on the chamber pressure transient. The analytical combustion model is verified by good correlation between predicted and observed chamber pressure transients. The developed model enables a rocket designer to vary the design parameters analytically to obtain stable combustion for a particular mode of operation which is prescribed by mission objectives.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

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

    SciTech Connect

    Guo, X.; Wehrmeyer, J.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-{var_epsilon} model, RNG k-{var_epsilon} 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. {copyright} {ital 1997 American Institute of Physics.}

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Deposit formation in hydrocarbon rocket fuels: Executive summary

    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 811K. In addition, the effect of the inside wall material on deposit formation was evaluated in selected tests which were conducted using nickel-plated tubes. The results of the tests indicated that substantial deposit formation occurs with RP-1 fuel at wall temperatures between 600 and 800K, with peak deposit formation occurring near 700K. No improvements were obtained when de-oxygenated 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. The 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, plating the inside wall of the tubes with nickel was found to significantly reduce carbon deposition rates for RP-1 fuel.

  8. 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

  9. Photosynthetic water splitting: A biotechnological approach to gaseous fuel synthesis

    SciTech Connect

    Greenbaum, E.; Reeves, M.

    1986-01-01

    Photosynthesis research in the Chemical Technology Division of Oak Ridge National Laboratory is focused on understanding the physicochemical aspects of photosynthesis, with specific application to developing a biotechnological process for production of gaseous fuels from renewable inorganic materials. This paper provides a general review of photosynthetic water splitting by intact microalgae and platinum-modified chloroplasts for the simultaneous photoproduction of molecular hydrogen and oxygen. 9 refs.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  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. 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.

  19. 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.

  20. 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.

  1. Gaseous fueled torch apparatus and fueling module therefor

    SciTech Connect

    Czerwinski, K.S.; Gabany, E.; Sharma, S.S.; Turko, J.W.

    1988-10-11

    This patent describes a fueling module for supplying natural gas to a natural gas fueled torch apparatus including a torch adapted for use in cutting or welding operations, the torch apparatus further including a source of oxygen for supplying oxygen to the torch, and the torch being selectively operable for combustion of a mixture of natural gas and oxygen, the fueling module connectable to an electric power source and being supplying natural gas to the torch apparatus at an elevated pressure from a relatively low pressure natural gas supply system, the fueling module comprising: fueling module inlet means connectable in fluid communication with the natural gas supply system; compression means in fluid communication with the fueling module inlet means and selectively energizable for compressing the natural gas from the natural gas supply system in order to increase its pressure, the compression means having a compression intake in fluid communication with the fueling module inlet means and a compression discharge outlet for discharging compressed natural gas from the compression means; lubricant filter means in fluid communication with the compression discharge outlet for substantially trapping and collecting compression means lubricants from the compressed natural gas from the compression discharge outlet and for returning the collected compression means lubricants to the compression intake; cooling means in fluid communication with the compression discharge outlet means for reducing the temperature of the compressed natural gas.

  2. 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.

  3. Gaseous fueled torch apparatus and fueling module therefor

    SciTech Connect

    Czerwinski, K.S.; Gabany, E.; Sharma, S.S.; Turko, J.W.

    1990-06-05

    This patent describes a fueling system. It is used for supplying natural gas to one of a natural gas fueled torch apparatus and at least one storage vessel, the torch apparatus including a torch adapted for use in cutting or welding operations. The torch apparatus includes a source of oxygen for supplying oxygen to the torch, and the torch being selectively operable for combustion of a mixture of natural gas and oxygen. The fueling system is connectable to an electric power source and supplies natural gas to one of the torch apparatus and the storage vessel at an elevated pressure from a relatively low pressure natural gas supply system.

  4. Gaseous fueled vehicles: A role for natural gas and hydrogen

    SciTech Connect

    Blazek, C.F.; Jasionowski, W.J.

    1991-01-01

    The commercialization of gaseous hydrogen fueled vehicles requires both the development of hydrogen fueled vehicles and the establishment of a hydrogen fueling infrastructure. These requirements create a classic chicken and egg scenario in that manufacturers will not build and consumers will not buy vehicles without an adequate refueling infrastructure and potential refueling station operators will not invest the needed capital without an adequate market to serve. One solution to this dilemma is to create a bridging strategy whereby hydrogen is introduced gradually via another carrier. The only contending alternative fuel that can act as a bridge to hydrogen fueled vehicles is natural gas. To explore this possibility, IGT is conducting emission tests on its dedicated natural gas vehicle (NGV) test platform to determine what, if any, effects small quantities of hydrogen have on emissions and performance. Furthermore, IGT is actively developing an adsorbent based low-pressure natural gas storage system for NGV applications. This system has also shown promise as a storage media for hydrogen. A discussion of our research results in this area will be presented. Finally, a review of IGT's testing facility will be presented to indicate our capabilities in conducted natural gas/hydrogen vehicle (NGHV) research. 3 refs., 10 figs.

  5. 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.

  6. Performance Characteristics of the Methane Fueled Rocket Nozzles

    NASA Astrophysics Data System (ADS)

    Ito, Takashi; Miyajima, Hiroshi

    Performance of the methane fueled rocket nozzles are numerically investigated using computational fluid dynamics approach. A simple set of chemical reactions and kinetics for methane/oxygen nozzle flow is proposed. The chamber pressure, mixture ratio and size of the nozzle are parametrically changed to study the influence of characteristic rocket engine design parameters on nozzle losses. The amount of dissociation is high when the chamber pressure is low and the kinetic loss becomes dominant compared to the other nozzle losses. The peak specific impulse is achieved at a higher mixture ratio region as the chamber pressure increases. The chemical non-equilibrium flow appears mainly at down stream region of the nozzle throat. The influence of the chemical non-equilibrium effect decreases as the chamber pressure increases. Supersonic chemically reactive gas stays longer in the nozzle as the size of the nozzle become larger and the amount of recombination increases which decreases the kinetic loss. When the chamber pressure is high, the kinetic loss becomes small and the effect of the size of nozzle also becomes small.

  7. 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.

  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. 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.

  10. 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.

  11. 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.

  12. 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

  13. 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

  14. Modeling of gaseous reacting flow and thermal environment of liquid rocket injectors

    NASA Astrophysics Data System (ADS)

    Sozer, Emre

    Reacting flow and thermal fields around the injector critically affect the performance and life of liquid rocket engines. The performance gain by enhanced mixing is often countered by increased heat flux to the chamber wall, which can result in material failure. A CFD based design approach can aid in optimization of competing objectives by providing detailed flow field data and an ability to feasibly evaluate a large number of design configurations. To address issues related to the CFD analysis of such flows, various turbulence and combustion modeling aspects are assessed. Laminar finite-rate chemistry and steady laminar flamelet combustion models are adopted to facilitate individual assessments of turbulence-chemistry interactions (TCI) and chemical non-equilibrium. Besides the experimental wall heat transfer information, assessments are aided by evaluations of time scales, grid sensitivity, wall treatments and kinetic schemes. Several multi-element injector configurations are considered to study element-to-element interactions. Under the conditions considered, chemical non-equilibrium effect is found to be unimportant. TCI is found to noticeably alter the flow and thermal fields near the injector and the flame surface. In the multi-element injector case, due to proximity of the outer row injector elements to the wall, wall heat flux distribution is also significantly affected by TCI. The near wall treatment is found to critically affect wall heat flux predictions. A zonal treatment, blending the low-Reynolds number model and the law-of-the-wall approach is shown to improve the accuracy significantly. Porous materials such as Rigimesh are often used as the injector face plate of liquid rocket engines. A multi-scale model which eliminates the empirical dependence of conventional analysis methods, is developed. The resulting model is tested using experimental information showing excellent agreement. The model development and assessment presented for both injector

  15. 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.

  16. Monomethylhydrazine versus hydrazine fuels - Test results using a 100 pound thrust bipropellant rocket engine

    NASA Technical Reports Server (NTRS)

    Smith, J. A.; Stechman, R. C.

    1981-01-01

    A test program was performed to evaluate hydrazine (N2H4) as a fuel for a 445 Newton (100 lbf) thrust bipropellant rocket engine. Results of testing with an identical thruster utilizing monomethylhydrazine (MMH) are included for comparison. Engine performance with hydrazine fuel was essentially identical to that experienced with monomethylhydrazine although higher combustor wall temperatures (approximately 400 F) were obtained with hydrazine. Results are presented which indicate that hydrazine as a fuel is compatible with Marquardt bipropellant rocket engines which use monomethylhydrazine as a baseline fuel.

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

    DOEpatents

    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.

  18. 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.

  19. 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.

  20. 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.

  1. 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

  2. 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)

  3. 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.

  4. 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.

  5. A study of mesospheric rocket contrails and clouds produced by liquid-fueled rockets

    NASA Technical Reports Server (NTRS)

    Turco, R. P.; Toon, O. B.; Whitten, R. C.; Keesee, R. G.; Hollenbach, D.

    1982-01-01

    Changes in the atmospheric composition, particularly through the condensation of rocket vehicle exhaust, caused by the flights of 400 heavy lift launch vehicles (HLLV) to carry crews and materials into space to build a satellite solar power system (SPS) were examined. Attention was given to the formation of mesospheric contrails and clouds. A one-dimensional model was used to formulate the photochemistry and vertical transport of water vapor, its nucleation into an ice cloud, and the microphysical development of the cloud. Considering one HLLV launch per day for a decade, it is projected that the upper atmosphere water vapor concentration would be increased by 10-20%, thereby augmenting the size and opacity of natural noctilucent clouds by 50%. No climatological consequences are foreseen from the clouds, although spectacular noctiluminescent cloud displays are thought to be possible.

  6. 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

    2016-02-01

    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. PMID:26687630

  7. Experimental investigation of laboratory-scale rocket engine fed on solid polyethylene rod as fuel

    NASA Astrophysics Data System (ADS)

    Yemets, V. V.; Sanin, F. P. Dzhur, Ye. O.; Masliany, M. V.; Kostritsyn, O. Yu.; Minteev, G. V.; Ushkanov, V. M.

    Fire testing of the laboratory-scale rocket engine with the consumable solid polyethylene rod as fuel is described. The experimental data on heat flows, gasification rate and heat transfer coefficient are presented. Results of the testing may be useful for designing launch vehicles with combustible polyethylene tank shells.

  8. Restart transients of hybrid rocket engines.

    NASA Technical Reports Server (NTRS)

    Saraniero, M. A.; Caveny, L. H.; Summerfield, M.

    1973-01-01

    Experimental investigation of the problems associated with restarting hybrid rocket motors (i.e., motors wherein a liquid or gaseous oxidizer is injected into the port of a solid fuel grain with subsequent mixing and combustion of the oxidizer and fuel) following a brief period of extinguishment. The results include the finding that the ignition delay on restart is decreased because less energy is absorbed by the fuel before the surface reaches the ignition point.

  9. Experimental investigation of fuel regression rate in a HTPB based lab-scale hybrid rocket motor

    NASA Astrophysics Data System (ADS)

    Li, Xintian; Tian, Hui; Yu, Nanjia; Cai, Guobiao

    2014-12-01

    The fuel regression rate is an important parameter in the design process of the hybrid rocket motor. Additives in the solid fuel may have influences on the fuel regression rate, which will affect the internal ballistics of the motor. A series of firing experiments have been conducted on lab-scale hybrid rocket motors with 98% hydrogen peroxide (H2O2) oxidizer and hydroxyl terminated polybutadiene (HTPB) based fuels in this paper. An innovative fuel regression rate analysis method is established to diminish the errors caused by start and tailing stages in a short time firing test. The effects of the metal Mg, Al, aromatic hydrocarbon anthracene (C14H10), and carbon black (C) on the fuel regression rate are investigated. The fuel regression rate formulas of different fuel components are fitted according to the experiment data. The results indicate that the influence of C14H10 on the fuel regression rate of HTPB is not evident. However, the metal additives in the HTPB fuel can increase the fuel regression rate significantly.

  10. 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

  11. 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.

  12. 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.

  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. 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.

  15. 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.

  16. 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

  17. 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.

  18. Supersonic-combustion rocket

    NASA Technical Reports Server (NTRS)

    Weber, R. J.; Franciscus, L. C. (Inventor)

    1973-01-01

    A supersonic combustion rocket is provided in which a small rocket motor is substituted for heavy turbo pumps in a conventional rocket engine. The substitution results in a substantial reduction in rocket engine weight. The flame emanating from the small rocket motor can act to ignite non-hypergolic fuels.

  19. 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.

  20. 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.

  1. 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

  2. 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.

  3. LOW NOX, HIGH EFFICIENCY MULTISTAGED BURNER: GASEOUS FUEL RESULTS

    EPA Science Inventory

    The paper discusses the evaluation of a multistaged combustion burner design on a 0.6 MW package boiler simulator for in-furnace NOx control and high combustion efficiency. Both deep air staging, resulting in a three-stage configuration, and boiler front wall fuel staging of undo...

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

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

    SciTech Connect

    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.

  10. 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.

  11. 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

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

    1992-09-01

    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.

  18. Hybrid rocket performance

    NASA Astrophysics Data System (ADS)

    Frederick, Robert A., Jr.

    1992-12-01

    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.

  19. 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.

  20. 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.

  1. 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.

  2. 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. PMID:23850117

  3. 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

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

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

    DOEpatents

    Wijmans Johannes G.; Merkel, Timothy C.; Baker, Richard W.

    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.

  10. 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.

  11. 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.

  12. 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

  13. Hybrid rocket motor testing at Nammo Raufoss A/S

    NASA Astrophysics Data System (ADS)

    Rønningen, Jan-Erik; Kubberud, Nils

    2005-08-01

    Hybrid rocket motor technology and the use of hybrid rockets have gained increased interest in recent years in many countries. A typical hybrid rocket consists of a tank containing the oxidizer in either liquid or gaseous state connected to the combustion chamber containing an injector, inert solid fuel grain and nozzle. Nammo Raufoss A/S has for almost 40 years designed and produced high-performance solid propellant rocket motors for many military missile systems as well as solid propellant rocket motors for civil space use. In 2003 an in-house technology program was initiated to investigate and study hybrid rocket technology. On 23 September 2004 the first in-house designed hybrid test rocket motor was static test fired at Nammo Raufoss Test Center. The oxidizer was gaseous oxygen contained in a tank pressurized to 10MPa, flow controlled through a sonic orifice into the combustion chamber containing a multi port radial injector and six bore cartridge-loaded fuel grain containing a modified HTPB fuel composition. The motor was ignited using a non-explosive heated wire. This paper will present what has been achieved at Nammo Raufoss since the start of the program.

  14. Numerical and experimental studies of the hybrid rocket motor with multi-port fuel grain

    NASA Astrophysics Data System (ADS)

    Tian, Hui; Li, Xintian; Zeng, Peng; Yu, Nanjia; Cai, Guobiao

    2014-03-01

    This paper presents three-dimensional numerical simulations and experimental studies of the hybrid rocket motor with multi-port fuel grain. The numerical model is established based on the Navier-Stokes equations with turbulence, chemical reactions, fuel pyrolysis, and solid-gas boundary interactions. The simulation is performed based on the 98% hydrogen peroxide (HP) and hydroxyl terminated polybutadiene (HTPB) propellant combination. The results indicate that the flow field and fuel regression rate distributions present apparent three-dimensional characteristics. The fuel regression rates decrease first and then gradually increase with the axial location increasing. At a certain cross section, the fuel regression rates are lower in the points on arcs with smaller radius of curvature when the fuel port is a derivable convex figure. Two experiments are carried out on a full scale motor with the simulation one. The working process of the motor is steady and no evident oscillatory combustion is observed. The fuel port profiles before and after tests indicate that the fuel regression rate distributions at the cross section match well with the numerical simulation results.

  15. 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. PMID:27270594

  16. Modeling of gaseous sup 14 CO sub 2 release from perforations in spent fuel disposal containers

    SciTech Connect

    Pescatore, C.; Sullivan, T.M.

    1991-11-01

    The potential release of gaseous {sup 14}CO{sub 2} from small perforations in spent fuel containers has been evaluated as a function of temperature, hole size, effective porosity of corrosion products within the hole, and time, based on the waste package design parameters and environmental conditions described in the Yucca Mountain Site Characterization Report (SCP). The SCP does not specify initial fill gas (argon) pressure and temperature. It is shown that, if significant {sup 14}C oxidation takes place during the initial, inert-gas phase, an incentive exists to initially underpressurize the containers. This will avoid large, spiked releases of gaseous {sup 14}CO{sub 2} and will result in delayed, smaller, and more uniform release rates over time. Therefore, larger size perforations could be tolerated while meeting the applicable regulations.

  17. Labscale testing techniques for hybrid rockets

    NASA Astrophysics Data System (ADS)

    Hollman, S. L.; Frederick, R. A., Jr.

    1993-06-01

    The objective of this work is to provide an overview of labscale testing techniques used for the testing of hybrid rocket motors. The hybrid rocket motor uses a solid fuel and a gaseous or liquid oxidizer. The review summarizes the important features of past U.S. testing, identifies key references, and assesses the relevance of the laboratory data to larger scale motors. The results show that important parameters can be observed through labscale testing of hybrid rockets. This may allow for proving theories and testing materials. The conclusions are that labscale testing is important in understanding the behavior of the hybrid rocket and in improving its performance, but additional work is needed in experiment design.

  18. 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.

  19. 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.

  20. 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.

  1. 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

  2. Liquefied Gaseous Fuels Spill Test Facility program: Eleven additional chemicals: Environmental Assessment

    SciTech Connect

    Not Available

    1989-12-01

    An Environmental Assessment (EA) has been prepared to assess the environmental consequences of spill testing eleven hazardous materials at the Liquefied Gaseous Fuels Spill Test Facility (LGFSTF) at Frenchman Flat, Nevada Test Site (NTS). These chemicals are: chlorosulfonic acid, fluorosulfonic acid, hydrogen chloride, methyl trichlorosilane, nitrogen tetroxide, oleum, silicon tetrachloride, sulfur-trioxide, titanium tetrachloride, trichlorosilane, and unsymmetrical dimethyl hydrazine. DOE has determined that the proposed spill testing of these eleven hazardous materials at LGFSTF at Frenchman Flat is not a major federal action significantly affecting the quality of the human environment within the meaning of the National Environmental Policy Act (NEPA). Therefore, an environmental impact statement (EIS) will not be prepared.

  3. 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,...

  4. 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

    ... 40 Protection of Environment 19 2010-07-01 2010-07-01 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,...

  5. 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.

  6. Hybrid Rocket Motor Test

    NASA Technical Reports Server (NTRS)

    1994-01-01

    A 10,000-pound thrust hybrid rocket motor is tested at Stennis Space Center's E-1 test facility. A hybrid rocket motor is a cross between a solid rocket and a liquid-fueled engine. It uses environmentally safe solid fuel and liquid oxygen.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. A Thrust and Impulse Study of Guanidinium Azo-Tetrazolate as an Additive for Hybrid Rocket Fuel

    NASA Astrophysics Data System (ADS)

    Patton, J.; Wright, A. M.; Dunn, L.; Alford, B.

    2000-03-01

    A thrust and impulse study of the hybrid rocket fuel additive Guanidinium Azo-Tetrazolate (GAT) was conducted at the University of Arkansas at Little Rock (UALR) Hybrid Rocket Facility. GAT is an organic salt with a high percentage of nitrogen. GAT was mixed with the standard hybrid rocket fuel, Hydroxyl-Terminated Polybutadiene (HTPB), in the concentration of 15%, by mass. The fuel grains with the GAT additive were fired for 4 second runs with the oxygen flows of 0.05, 0.07, 0.09, and 0.12 lbm/sec. For each run average thrust, total impulse, and specific impulse were measured. Average thrust, specific impulse, and total impulse vs. oxygen flow were plotted. Similar data was collected for plain HTPB/PAPI fuels for comparison. GAT was found to increase the thrust output when it was added to the standard hybrid rocket fuel, HTPB. GAT also increased the total impulse during the run. The thrust and total impulse were increased at all flows, but especially at the lower oxygen flow rates. Specific impulse only increased during the lower oxygen flow runs, and decreased slightly for the higher oxygen flow runs.

  12. 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.

  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. A physics-based two-dimensional comprehensive mathematical model to predict non-uniform regression rate in solid fuels for hybrid rocket motors

    NASA Astrophysics Data System (ADS)

    Antoniou, Antonis

    A numerical study using a comprehensive physics based mathematical model is conducted to predict the fuel regression rate in hybrid rocket fuels. The physical model adopted for the study is based on an unsteady, two-domain (solid fuel and gaseous oxidizer coupled through a moving interface) concept where both domains are assumed to be two-dimensional. The oxidizer gas flow is assumed to be compressible and turbulent with Navier-Stokes Assumptions. The radiative heat transfer is incorporated to the energy equation for the gas domain using the Rosseland diffusion approximation. Fuel is assumed to be a nontransparent isotropic solid. The two domains are coupled through an energy balance at the interface that includes heat transfer due to radiation, conduction, and ablation. The regression rate of the fuel surface due to ablation is modeled using the first-order Arrhenius Equation. The combustion of the ablated fuel is modeled by single step, three species chemical reaction equation of second order Arrhenius type. The solution to the governing differential equations of the present model is obtained by first transform the solution domain using a time and space dependent transformation. In the gas domain the transformed set of differential equations is discretized by a fully implicit finite-difference technique then linearized by using Newton linearization method. The resulting set of algebraic equations are transformed by the Coupled Modified Strongly Implicit Procedure (CMSIP) for the primitive variables of the problem. Validation of the solution algorithm and the CMSIP that is developed for this study is validated through the study of two bench mark cases: driven cavity and flow through channel. Furthermore, the results of the comprehensive model are compared to those of the parabolic incompressible model. Finally the proposed comprehensive mathematical model is used to predict the unsteady temperature and pressure distributions, and the velocity field in the gas

  15. 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.

  16. 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

  17. Experimental investigation of a solid rocket combustion simulator

    NASA Technical Reports Server (NTRS)

    Frederick, Robert A., Jr.

    1991-01-01

    The response of solid rocket motor materials to high-temperature corrosive gases is usually accomplished by testing the materials in a subscale solid rocket motor. While this imposes the proper thermal and chemical environment, a solid rocket motor does not provide practical features that would enhance systematic evaluations such as: the ability to throttle for margin testing, on/off capability, low test cost, and a low-hazards test article. Solid Rocket Combustion Simulators (SRCS) are being evaluated by NASA to test solid rocket nozzle materials and incorporate these essential practical features into the testing of rocket materials. The SRCS is designed to generate the thermochemical environment of a solid rocket. It uses hybrid rocket motor technology in which gaseous oxygen (Gox) is injected into a chamber containing a solid fuel grain. Specific chemicals are injected in the aft mixing chamber so that the gases entering the test section match the temperature and a non-dimensional erosion factor B' to insure similarity with a solid motor. Because the oxygen flow can be controlled, this approach allows margin testing, the ability to throttle, and an on/off capability. The fuel grains are inert which makes the test article very safe to handle. The objective of this work was to establish the baseline operating characteristics of a Labscale Solid Rocket Combustion Simulator (LSRCS). This included establishing the baseline burning rates of plexiglass fuels and the evaluation of a combustion instability for hydroxy-terminated polybutadyene (HTPB) propellants. The scope of the project included: (1) activation of MSFC Labscale Hybrid Combustion Simulator; (2) testing of plexiglass fuel at Gox ranges from 0.025 to 0.200 lb/s; (3) burning HTPB fuels at a Gox rate of 0.200 lb/s using four different mixing chamber configurations; and (4) evaluating the fuel regression and chamber pressure responses of each firing.

  18. 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.

  19. 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%. PMID:25088312

  20. Predicting gaseous emissions from small-scale combustion of agricultural biomass fuels.

    PubMed

    Fournel, S; Marcos, B; Godbout, S; Heitz, M

    2015-03-01

    A prediction model of gaseous emissions (CO, CO2, NOx, SO2 and HCl) from small-scale combustion of agricultural biomass fuels was developed in order to rapidly assess their potential to be burned in accordance to current environmental threshold values. The model was established based on calculation of thermodynamic equilibrium of reactive multicomponent systems using Gibbs free energy minimization. Since this method has been widely used to estimate the composition of the syngas from wood gasification, the model was first validated by comparing its prediction results with those of similar models from the literature. The model was then used to evaluate the main gas emissions from the combustion of four dedicated energy crops (short-rotation willow, reed canary grass, switchgrass and miscanthus) previously burned in a 29-kW boiler. The prediction values revealed good agreement with the experimental results. The model was particularly effective in estimating the influence of harvest season on SO2 emissions. PMID:25543541

  1. 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.

  2. 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.

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

    SciTech Connect

    Mahalatkar, K.; Kuhlman, J.; Huckaby, E.D.; O'Brien, T.

    2011-01-01

    A computational fluid dynamic(CFD) model for the fuel reactor of chemical looping combustion technology has been developed,withspecialfocusonaccuratelyrepresentingtheheterogeneous chemicalreactions.Acontinuumtwo-fluidmodelwasusedtodescribeboththegasandsolidphases. Detailedsub-modelstoaccountforfluid–particleandparticle–particleinteractionforceswerealso incorporated.Twoexperimentalcaseswereanalyzedinthisstudy(Son andKim,2006; Mattisonetal., 2001). SimulationswerecarriedouttotestthecapabilityoftheCFDmodeltocapturechangesinoutletgas concentrationswithchangesinnumberofparameterssuchassuperficialvelocity,metaloxide concentration,reactortemperature,etc.Fortheexperimentsof Mattissonetal.(2001), detailedtime varyingoutletconcentrationvalueswerecompared,anditwasfoundthatCFDsimulationsprovideda reasonablematchwiththisdata.

  4. 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.

  5. 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

  6. 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.

  7. 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.

  8. 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

  9. 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.

  10. MODIFICATION OF SPILL FACTORS AFFECTING AIR POLLUTION. VOLUME II. THE CONTROL OF THE VAPOR HAZARD FROM SPILLS OF LIQUID ROCKET FUELS

    EPA Science Inventory

    The hypergolic rocket fuels, hydrazine and nitrogen tetroxide, are volatile hazardous materials of special interest to the Air Force. Through monitoring of ongoing Environmental Protection Agency programs, the Air Force has maintained cognizance of the developing state of the art...

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

    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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. EPA evaluation of the VCD supplemental gaseous fuel delivery system under section 511 of the Motor Vehicle Information and Cost Savings Act. Technical report

    SciTech Connect

    Barth, E.A.

    1983-09-01

    This report announces the conclusions of the Environmental Protection Agency (EPA) evaluation of the 'VCD Supplemental Gaseous Fuel Delivery System' under the provisions of Section 511 of the Motor Vehicle Information and Cost Savings Act. The evaluation of the 'VCD Supplemental Gaseous Fuel Delivery System' was conducted on the application of the manufacturer. The device is designed to operate the engine of a vehicle on a mixture of gasoline and propane. The device consists of a gaseous fuel metering and control unit, a modified carburetor and associated electrical and plumbing components. It functions by replacing some of the gasoline with propane under certain operating conditions. The device causes the engine to idle on propane, cruise on gasoline, and accelerate on a mixture of the two fuels. This is claimed to be more fuel efficient. This combination of improvements in fuel efficency and fuel substitution is claimed to save both fuel and money.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Simulation of the internal dynamics of solid-fuel rocket engines on the basis of the STAR-CD suite

    NASA Astrophysics Data System (ADS)

    Volkov, K. N.; Denisikhin, S. V.; Emel'Yanov, V. N.

    2006-07-01

    Flows in the channels of solid-fuel charges with cross sections having different views in plan and flows in the prenozzle volume and the nozzle unit of a solid-fuel rocket engine have been simulated on the basis of the STAR-CD suite for different types of charges and different designs of the input part of the engine nozzle. The influence of the compressibility, turbulence, geometric factors, and flow rate on the distributions of gasdynamic parameters in the working region of the engine has been investigated.

  2. 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.

  3. Regression rate and pyrolysis behavior of HTPB-based solid fuels in a hybrid rocket motor

    NASA Astrophysics Data System (ADS)

    Chiaverini, Martin John

    An experimental investigation on the regression rate and pyrolysis behavior of hydroxyl-terminated polybutadiene-based solid fuels has been conducted. The overall objective was to obtain a better understanding of the physical processes governing solid-fuel regression and pyrolysis under different operating regimes. Experiments were conducted using a windowed, slab geometry hybrid motor and a conductive-heating induced thermal pyrolysis test rig. Gaseous oxygen was employed as the oxidizer in the 1-m long, lab-scale hybrid motor, which had realistic operating conditions. A real-time X-ray radiography system and an ultrasonic pulse-echo system were both used to obtain the local, instantaneous solid fuel regression rates. A semi-empirical approach was developed to analyze the experimental results and to correlate the regression rates with physically descriptive, dimensionless parameters. For relatively high surface temperatures above 722 K, the activation energy of pure HTPB was 4.91 kcal/mole, indicating that the pyrolysis process was governed by formation and desorption of high molecular weight fragments from the fuel surface. The conductive-heating induced pyrolysis rates of HTPB, conducted at atmospheric pressure, were very similar to those measured in the hybrid motor tests at much higher pressures. This result implies that the regression rate of HTPB was governed primarily by thermal decomposition processes and not influenced by heterogeneous surface reactions. Radiant heat transfer had a significant effect on the overall regression rate behavior of HTPB. Radiation from soot generally accounted for about 80 to 90% of the total radiant heat flux. Two separate expressions, one for the developing flow regime and one for fully-developed flow, were used to correlate the regression rate data. Both correlations show that standard hybrid boundary layer correlations must be modified to account for the effects of variable fluid properties across the boundary layer and

  4. Combustion Characteristics of a Swirling LOX Type Hybrid Rocket Engine

    NASA Astrophysics Data System (ADS)

    Kitagawa, Koki; Yuasa, Saburo

    We have proposed a swirling oxidizer type hybrid rocket engine. In this paper, liquid oxygen (LOX) was used as oxidizer. Combustion tests of a hybrid rocket engine with a swirling LOX flow were conducted by changing the swirl strength. Ignition was rapid and reliable, and combustion of PMMA with swirling LOX was stable. Fuel regression rates, C* efficiency and specific impulse of the hybrid rocket engine with swirling LOX flow were smaller than those with swirling gaseous oxygen (GOX). This low performance may be restraint of atomization and vaporization of LOX by formation of a liquid layer on the PMMA fuel and a decline of angular momentum of the swirling LOX during vaporization. Combustion oscillation occurred when the ratios of differential pressure between injector pressure and chamber pressure to chamber pressure were small. This combustion oscillation was confirmed to be a “Chugging” mode due to combustion time lag of LOX.

  5. 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

  6. 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.

  7. 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.

  8. Modeling of gaseous {sup 14}CO{sub 2} release from perforations in spent fuel disposal containers

    SciTech Connect

    Pescatore, C.; Sullivan, T.M.

    1991-11-01

    The potential release of gaseous {sup 14}CO{sub 2} from small perforations in spent fuel containers has been evaluated as a function of temperature, hole size, effective porosity of corrosion products within the hole, and time, based on the waste package design parameters and environmental conditions described in the Yucca Mountain Site Characterization Report (SCP). The SCP does not specify initial fill gas (argon) pressure and temperature. It is shown that, if significant {sup 14}C oxidation takes place during the initial, inert-gas phase, an incentive exists to initially underpressurize the containers. This will avoid large, spiked releases of gaseous {sup 14}CO{sub 2} and will result in delayed, smaller, and more uniform release rates over time. Therefore, larger size perforations could be tolerated while meeting the applicable regulations.

  9. 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.

  10. 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.

  11. 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. PMID:21081245

  12. Global stratospheric effects of the alumina emissions by solid-fueled rocket motors

    NASA Astrophysics Data System (ADS)

    Danilin, M. Y.; Shia, R.-L.; Ko, M. K. W.; Weisenstein, D. K.; Sze, N. D.; Lamb, J. J.; Smith, T. W.; Lohn, P. D.; Prather, M. J.

    2001-01-01

    We simulate accumulation of Al2O3 particles in the atmosphere produced by solid-fueled rocket motors by using the Goddard Institute for Space Studies/University of California at Irvine three-dimensional (3-D) chemistry-transport model (CTM). Our study differs from Jackman et al. (1998) by applying a 3-D CTM, considering 13 size bins for the emitted particles from 0.025 to 10 μm and taking into account their washout, gravitational sedimentation, and coagulation with background sulfate aerosol. We assume an initial trimodal size distribution of Al2O3 particles (Beiting, 1997) with 2.8% by mass of the alumina emitted as particles with radius of less than 1 μm. Our test case adopts a stratospheric source of 1120 tons/yr equivalent to nine space Shuttle and four Titan IV launches annually. The calculated steady state surface area density (SAD) and mass density for the scenarios with sedimentation of alumina particles have maximum values in the lower stratosphere in the Northern Hemisphere of up to 7×10-4 μm2/cm3 and 0.09 ng/m3, respectively, or about 1000 times smaller than those of the background sulfate aerosol. Our results are sensitive to the emitted mass fractionation of alumina (EMFA) showing the values for the SAD or mass density higher or lower by an order of magnitude owing to a poorly known EMFA. Chemical implications of alumina particle accumulation for the ozone balance are estimated by using the Atmospheric and Environmental Research 2-D model assuming chlorine activation on Al2O3 surfaces via the C1ONO2 + HCl → Cl2 + HNO3 reaction with a probability of 0.02 (Molina et al., 1997). Owing to the very small Al2O3 SAD, any additional ozone depletion due to Al2O3 emissions is also small (0.0028% on a global annually averaged basis for the scenario with sedimentation, or about 4 times smaller than the ozone response to chlorine emissions only). The ozone depletion potential of the alumina emissions is about 0.03-0.08 for the scenarios using the EMFA of

  13. 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.

  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. 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)

  16. 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

  17. 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

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Analysis and testing of similarity and scale effects in hybrid rocket motors

    NASA Astrophysics Data System (ADS)

    Dayal Swami, Rajeshwar; Gany, Alon

    2003-04-01

    In order to derive proper scaling rules in hybrid rocket motors, a theoretical similarity analysis is presented. By taking account of the main phenomena and effects, the similarity analysis defines the following three main conditions for testing a laboratory-scale hybrid rocket motor that can simulate a full-scale motor: (1) geometric similarity, (2) same fuel and oxidizer combination, and (3) scaling mass flow rate of oxidizer in proportion to the motor port diameter. To verify the analysis, tests are conducted on different-size polymethylmethacrylate/gaseous oxygen hybrid rocket motors. These motors are scaled as per the similarity analysis and tested under similarity conditions. A fairly good agreement between the test-results and theoretical prediction verifies the similarity model. This also points out that the main processes and effects associated with hybrid rocket combustion have been considered adequately in the analysis.

  5. 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.

  6. Rocket motor vulnerability considerations in relation to bullet impact and fuel fires

    NASA Astrophysics Data System (ADS)

    Mason, A. C.

    1992-07-01

    This paper reports on the work undertaken by Royal Ordnance in relation to the assessment of solid propellant rocket motor vulnerability. A general overview of the RO IM Database is included with specific details being given on the results of 183 half inch bullet impact and 43 fast cook-off tests. The large number of trials conducted allows some statistical conclusions to be drawn and these can be used to design motors having a high probability of meeting the bullet impact and fast cook-off requirement of MIL-STD-2105 and STANAG 4241/STANAG 4240.

  7. 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.

  8. Heavy duty liquid and gaseous fuel emissions database test results from four alternative fuel configurations of the Caterpillar 3406 engine

    SciTech Connect

    Waldman, D.J. )

    1990-06-01

    Through the cooperation of several organizations including the Oak Ridge National Laboratory (ORNL) acting under the auspices of the Doe Alternative Fuels Utilization Program, heavy duty transient and steady-state emissions tests were conducted on four alternative fuel configurations of the Caterpillar 3406 engine. These included a diesel baseline, glow plug ignited methanol (diesel cycle), lean-burn spark ignited natural gas, and dual fuel (diesel pilot ignited natural gas). Results indicated methanol and natural gas both show excellent potential for low NOx and low particulate emissions. With these fuels however, unburned fuel emissions were much higher, especially in the dual fuel case, than the diesel baseline. Particulate emissions from the methanol and lean burn gas engines are thought to be almost entirely lube oil sourced. All of the configurations will require significant reduction in hydrocarbon and/or particulate emissions in order to meet the 1994 EPA emissions standards for heavy duty truck engines. 3 refs., 23 figs., 15 tabs.

  9. 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.

  10. 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.

  11. 40 CFR 86.1310-2007 - Exhaust gas sampling and analytical system for gaseous emissions from heavy-duty diesel-fueled...

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 19 2010-07-01 2010-07-01 false Exhaust gas sampling and analytical system for gaseous emissions from heavy-duty diesel-fueled engines and particulate emissions from all engines. 86.1310-2007 Section 86.1310-2007 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL...

  12. 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.

  13. 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.

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

    NASA Technical Reports Server (NTRS)

    Lohr, Jonathan C.; Trinh, Huu P.

    1992-01-01

    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.

  15. 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.

  16. 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.

  17. 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.

  18. 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. PMID:25947054

  19. 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.

  20. Partitioning of metal species during an enriched fuel combustion experiment. speciation in the gaseous and particulate phases.

    PubMed

    Pavageau, Marie-Pierre; Morin, Anne; Seby, Fabienne; Guimon, Claude; Krupp, Eva; Pécheyran, Christophe; Poulleau, Jean; Donard, Olivier F X

    2004-04-01

    Combustion processes are the most important source of metal in the atmosphere and need to be better understood to improve flue gas treatment and health impact studies. This combustion experiment was designed to study metal partitioning and metal speciation in the gaseous and particulate phases. A light fuel oil was enriched with 15 organometallic compounds of the following elements: Pb, Hg, As, Cu, Zn, Cd, Se, Sn, Mn, V, Tl, Ni, Co, Cr, and Sb. The resulting mixture was burnt in a pilot-scale fuel combustion boiler under controlled conditions. After filtration of the particles, the gaseous species were sampled in the stack through a heated sampling tube simultaneously by standardized washing bottles-based sampling techniques and cryogenically. The cryogenic samples were collected at -80 degrees C for further speciation analysis by LT/GC-ICPMS. Three species of selenium and two of mercury were evidenced as volatile species in the flue gas. Thermodynamic predictions and experiments suggest the following volatile metal species to be present in the flue gas: H2Se, CSSe, CSe2, SeCl2, Hg(0), and HgCl2. Quantification of volatile metal species in comparison between cryogenic techniques and the washing bottles-based sampling method is also discussed. Concerning metal partitioning, the results indicated that under these conditions, at least 60% (by weight) of the elements Pb, Sn, Cu, Co, Tl, Mn, V, Cr, Ni, Zn, Cd, and Sb mixed to the fuel were found in the particulate matter. For As and Se, 37 and 17%, respectively, were detected in the particles, and no particulate mercury was found. Direct metal speciation in particles was performed by XPS allowing the determination of the oxidation state of the following elements: Sb(V), Tl(III), Mn(IV), Cd(II), Zn(II), Cr(III), Ni(II), Co(II), V(V), and Cu(II). Water soluble species of inorganic Cr, As, and Se in particulate matter were determined by HPLC/ICP-MS and identified in the oxidation state Cr(III), As(V), and Se(IV). PMID

  1. 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.

  2. 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.

  3. 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.

  4. [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. PMID:27266032

  5. 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.

  6. 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.

  7. 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.

  8. 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)

  9. 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

  10. 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.

  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. 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.

  13. 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.

  14. 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

  15. Low thrust chemical rocket technology

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    1992-01-01

    A technology program aimed at improving the performance of low thrust chemical rockets for spacecraft onboard applications is reviewed. Navier-Stokes analyses of low Reynolds number rocket flows have been compared with local flow property measurements obtained using Rayleigh and Raman diagnostics in a 100 N gaseous hydrogen/gaseous oxygen rocket. It is indicated that computational domain should include the near injector flow and that the shear layer combustion model needs improvement. The system analyses and technical efforts intended to develop a technology base for higher performance propellants are presented. A LOX/hydrazine engine is demonstrated to have a 95 percent theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1.

  16. 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

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Laminar flow instability in nuclear rockets

    SciTech Connect

    Black, D.L. )

    1993-01-20

    Laminar flow instability (LFI) is a rarely encountered phenomenon, occurring in gaseous heated channels with high exit-to-inlet temperature ratios and a laminar Reynolds Number at the channel exit, as may be experienced in a nuclear rocket. Analytical techniques were developed and programmed for parametric evaluation that had been previously validated by comparison with available experimental data. The four types of transients associated with LFI are described in terms of the governing equations. Parametric evaluations of solid core prismatic and particle bed fuel configurations were made to determine their sensitivities to LFI from temperature ratio, flow rate, orificing, transition Reynolds Number, pressure level, presence of an exit sonic nozzle, power density and heat flux shape. The flow rate at the point of neutral stability and the growth rate of the excursive transient are calculated. The full power design point and the cooldown phases of operation were both evaluated.

  2. 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.

  3. 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.

  4. 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.

  5. 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%.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. Transient combustion in hybrid rockets

    NASA Astrophysics Data System (ADS)

    Karabeyoglu, Mustafa Arif

    1998-09-01

    Hybrid rockets regained interest recently as an alternative chemical propulsion system due to their advantages over the solid and liquid systems that are currently in use. Development efforts on hybrids revealed two important problem areas: (1) low frequency instabilities and (2) slow transient response. Both of these are closely related to the transient behavior which is a poorly understood aspect of hybrid operation. This thesis is mainly involved with a theoretical study of transient combustion in hybrid rockets. We follow the methodology of identifying and modeling the subsystems of the motor such as the thermal lags in the solid, boundary layer combustion and chamber gasdynamics from a dynamic point of view. We begin with the thermal lag in the solid which yield the regression rate for any given wall heat flux variation. Interesting phenomena such as overshooting during throttling and the amplification and phase lead regions in the frequency domain are discovered. Later we develop a quasi-steady transient hybrid combustion model supported with time delays for the boundary layer processes. This is integrated with the thermal lag system to obtain the thermal combustion (TC) coupled response. The TC coupled system with positive delays generated low frequency instabilities. The scaling of the instabilities are in good agreement with actual motor test data. Finally, we formulate a gasdynamic model for the hybrid chamber which successfully resolves the filling/emptying and longitudinal acoustic behavior of the motor. The TC coupled system is later integrated to the gasdynamic model to obtain the overall response (TCG coupled system) of gaseous oxidizer motors with stiff feed systems. Low frequency instabilities were also encountered for the TCG coupled system. Apart from the transient investigations, the regression rate behavior of liquefying hybrid propellants such as solid cryogenic materials are also studied. The theory is based on the possibility of enhancement

  11. Rockets using Liquid Oxygen

    NASA Technical Reports Server (NTRS)

    Busemann, Adolf

    1947-01-01

    It is my task to discuss rocket propulsion using liquid oxygen and my treatment must be highly condensed for the ideas and experiments pertaining to this classic type of rocket are so numerous that one could occupy a whole morning with a detailed presentation. First, with regard to oxygen itself as compared with competing oxygen carriers, it is known that the liquid state of oxygen, in spite of the low boiling point, is more advantageous than the gaseous form of oxygen in pressure tanks, therefore only liquid oxygen need be compared with the oxygen carriers. The advantages of liquid oxygen are absolute purity and unlimited availability at relatively small cost in energy. The disadvantages are those arising from the impossibility of absolute isolation from heat; consequently, allowance must always be made for a certain degree of vaporization and only vented vessels can be used for storage and transportation. This necessity alone eliminates many fields of application, for example, at the front lines. In addition, liquid oxygen has a lower specific weight than other oxygen carriers, therefore many accessories become relatively larger and heavier in the case of an oxygen rocket, for example, the supply tanks and the pumps. The advantages thus become effective only in those cases where definitely scheduled operation and a large ground organization are possible and when the flight requires a great concentration of energy relative to weight. With the aim of brevity, a diagram of an oxygen rocket will be presented and the problem of various component parts that receive particularly thorough investigation in this classic case but which are also often applicable to other rocket types will be referred to.

  12. Gaseous hydrogen leakage optical fibre detection system

    NASA Astrophysics Data System (ADS)

    Trouillet, Alain; Veillas, Colette; Sigronde, E.; Gagnaire, Henri; Clement, Michel

    2004-06-01

    Liquid hydrogen has been intensively used in aerospace applications during the past forty years and is of great interest for fuel cells technologies and future automotive applications. Following upon major explosive risks due to the use of hydrogen in air, previous studies were carried out in our laboratory in order to develop optical fiber sensors for the detection of hydrogen leakage. This communication is aimed towards a prototype optical fiber system designed for the detection of gaseous hydrogen leakage near the conecting flanges of the liquid hydrogen pipes on the test bench of the engine Vulcain of the rocket ARIANE V. Depending on the configuration, the prototype sensor provides a two-level alarm signal and the detection of gaseous hydrogen leakage is possible for concentrations lower than the lower explosive limit in air (between 0.1 and 4%) with alarm response times lower than 10 seconds in a wide range of temperatures between -35°C and 300°C. The sensing principle based on palladium-hydrogen interaction is presented as well as the detection system composed of an optical fiber probe and an optoelectronic device.

  13. Experimental and analytical comparison of flowfields in a 110 N (25 lbf) H2/O2 rocket

    NASA Technical Reports Server (NTRS)

    Reed, Brian D.; Penko, Paul F.; Schneider, Steven J.; Kim, Suk C.

    1991-01-01

    A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket was examined through the RPLUS code using the full Navier-Stokes equations with finite rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses were performed for a range of mixture ratios and fuel film cooling pcts. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four pct. Thrust coefficients computed by the code are found to be within two pct. of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty.

  14. 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.

  15. On-line molecular iodine isotopologue detection in gaseous media during spent nuclear fuel reprocessing using a laser-induced fluorescence method

    NASA Astrophysics Data System (ADS)

    Kireev, S. V.; Shnyrev, S. L.

    2015-06-01

    The paper reports on on-line measurement of the {}129{{\\text{I}}2}, 127I129I, and {}127{{\\text{I}}2} concentrations during spent nuclear fuel (SNF) reprocessing using a laser-induced fluorescence method. A He-Ne laser (632.8 nm) was used as a fluorescence excitation source. The detection limits obtained for molecular iodine isotopologue concentrations demonstrate the possibility of using this method for iodine control both in gaseous technological media generated during SNF reprocessing and after passing through the gas purification system (in atmosphere emission).

  16. 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

  17. 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.

  18. Computational fluid dynamic analysis of hybrid rocket combustor flowfields

    NASA Technical Reports Server (NTRS)

    Venkateswaran, S.; Merkle, C. L.

    1995-01-01

    Computational fluid dynamic analyses of the Navier-Stokes equations coupled with solid-phase pyrolysis, gas-phase combustion, turbulence and radiation are performed to study hybrid rocket combustor flowfields. The computational study is closely co-ordinated with a companion experimental program using a planar slab burner configuration with HTPB as fuel and gaseous oxygen. Computational predictions agree reasonably well with measurement data of fuel regression rates and surface temperatures. Additionally, most of the parametric trends predicted by the model are in general agreement with experimental trends. The computational model is applied to extend the results from the lab-scale to a full-scale axisymmetric configuration. The numerical predictions indicate that the full-scale configuration burns at a slower rate than the lab-scale combustor under identical specific flow rate conditions. The results demonstrate that detailed CFD analyses can play a useful role in the design of hybrid combustors.

  19. 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)

  20. Dumbo: A pachydermal rocket motor

    NASA Technical Reports Server (NTRS)

    Kirk, Bill

    1991-01-01

    A brief historical account is given of the Dumbo nuclear reactor, a type of folded flow reactor that could be used for rocket propulsion. Much of the information is given in viewgraph form. Viewgraphs show details of the reactor system, fuel geometry, and key characteristics of the system (folded flow, use of fuel washers, large flow area, small fuel volume, hybrid modulator, and cermet fuel).

  1. A six degree-of-freedom thrust sensor for a labscale hybrid rocket

    NASA Astrophysics Data System (ADS)

    Wright, Ann M.; Wright, Andrew B.; Born, Traig; Strickland, Ryan

    2013-12-01

    A six degree-of-freedom thrust sensor was designed, constructed, calibrated, and tested using the labscale hybrid rocket at the University of Arkansas at Little Rock. The system consisted of six independent legs: one parallel to the axis of symmetry of the rocket for main thrust measurement, two vertical legs near the nozzle end of the rocket, one vertical leg near the oxygen input end of the rocket, and two separated horizontal legs near the nozzle end. Each leg was composed of a rotational bearing, a load cell, and a universal joint above and below the load cell. The leg was designed to create point contact along only one direction and minimize the non-axial forces applied to the load cell. With this system, force in each direction and moments for roll, pitch, and yaw can be measured. The system was calibrated and tested using a labscale hybrid rocket using gaseous oxygen and hydroxyl-terminated polybutadiene solid fuel. The thrust stand proved to be stable during calibration tests. Thrust force vector components and roll, pitch, and yaw moments were calculated for test firings with an oxygen mass flow rate range of 0.0174-0.0348 kg s-1.

  2. 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.

  3. 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.

  4. Regression Rate Enhancement of Hybrid Rocket Motors using Mixed Hybrid Concept

    NASA Astrophysics Data System (ADS)

    Chidambaram, Palani Kumar; Kumar, Amit

    2011-11-01

    Low regression rates have been a major problem for hybrid rocket motors. In the present study, the effect on regression rate by adding ammonium perchlorate (AP) in solid fuel is studied numerically. AP mixed with HTPB is used as solid fuel and gaseous oxygen (GOX) is used as oxidizer. Solid fuel compositions are chosen such that the rocket motor retains start-stop capability. A reduced three step mechanism proposed in the literature is utilized to simulate the combustion. In the combustion chamber, two distinct flame fronts are captured. AP decomposition reaction forms a premixed flame front near the fuel surface. The AP decomposed products also react with HTPB. Heat released in these reactions improves the heat transferred to solid fuel and the regression rate significantly. Un-burnt fuel in the products further reacts with GOX forming a diffusion flame front farther from fuel surface. The presence of premixed flame front thus overcomes the low-regressing nature of hybrid combustion. It is found that 50% AP in solid fuel increases the regression rate by as much as 3 times.

  5. Nuclear rocket plume studies

    NASA Astrophysics Data System (ADS)

    Hastings, Daniel

    1993-05-01

    A description and detailed computational analysis of a vortex cleaning system designed to remove radioactive material from the plumes of nuclear rockets is included. The proposed system is designed to remove both particulates and radioactive gaseous material from the plume. A two part computational model is used to examine the system's ability to remove particulates, and the results indicate that under some conditions, the system can remove over 99% of the particles in the flow. Two critical parameters which govern the effectiveness of the system are identified and the information necessary to estimate cleaning efficiencies for particles of known sizes and densities is provided. A simple steady analytical solution is also developed to examine the system's ability to remove gaseous radioactive material. This analysis, while inconclusive, suggests that the swirl rates necessary to achieve useful efficiencies are too high to be achieved in any practical manner. Therefore, this system is probably not suitable for use, with gaseous radioactive material. It was concluded that the system can cause negligible specific impulse losses, though there may be a substantial mass penalty associated with its use.

  6. 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)

  7. 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.

  8. 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.

  9. 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.

  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. 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.

  12. 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.

  13. 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.

  14. Sirius-5 experimental rocket

    NASA Astrophysics Data System (ADS)

    Kerstein, A.; Omersel, P.; Goljuf, L.; Zidaric, M.

    1981-09-01

    After giving a historical account of multistage rocket development in Yugoslavia, a status report is presented for the three-stage Sirius-5 program. The rocket is composed of: (1) a solid-propellant first stage, consisting of a cluster of eight standard motors yielding 220 kN thrust for 1.3 sec; (2) a mixed amines/inhibited red fuming nitric acid, bipropellant second stage generating 50 kN thrust; and (3) a third stage of the same design as the second but with only 62 kg of fuel, by contrast to 168 kg. Among the design principles adhered to are: minimization of the number of components, conservative design margins, and specifications for key subsystems based on demonstration programs. The primary use of this system is in amateur rocketry, being able to carry a 20 kg payload to 150 km.

  15. Advanced rocket propulsion

    NASA Astrophysics Data System (ADS)

    Obrien, Charles J.

    1993-02-01

    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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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

    2012-10-01

    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. PMID:22913288

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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

  6. 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

  7. 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.

  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. 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)

  10. 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.

  11. 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).

  12. Hybrid Rocket Propulsion for Sounding Rocket Applications

    NASA Technical Reports Server (NTRS)

    1991-01-01

    A discussion of the H-225K hybrid rocket motor, produced by the American Rocket Company, is given. The H-225K motor is presented in terms of the following topics: (1) hybrid rocket fundamentals; (2) hybrid characteristics; and (3) hybrid advantages.

  13. 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. PMID:27219742

  14. 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. PMID:26247801

  15. Preliminary investigations on improving air-augmented rocket performance

    NASA Astrophysics Data System (ADS)

    Anil, K. N.; Damodaran, K. A.

    1994-05-01

    Use of the Petal nozzle instead of the conventional conical nozzle as the primary stream representing fuel-rich gases exiting from a rocket nozzle has demonstrated considerable improvement in the performance of an air-augmented rocket. This can be attributed to the improved mixing of the hot, exhaust gases containing unburnt fuel with the surrounding airstream, and subsequent heat release.

  16. 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.

  17. 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.

  18. 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.

  19. 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).

  20. 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.

  1. 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.

  2. Detection of Metallic Compounds in Rocket Plumes

    NASA Astrophysics Data System (ADS)

    Rogers, Chris; Dunn, Dr. Robert

    1998-04-01

    Recent experiments using metal mixed in hydroxyl-terminated polybutadiene (HTPB) fuel grains in small hybrid rocket indicates ion detectors may be effective in detection of metallic compounds in rocket plumes. We wanted to ascertain the extent to which the presence of metallic compounds in rocket plumes could be detected using ion probes and Gaussian rings. Charges that collide with or pass near the intruding probe are detected. Gaussian rings, short insulated cylindrical Gaussian surfaces, enclose the plume without intruding into the plume. Charges in the plume are detected by currents they induce in the cylinder.

  3. 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

  4. 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.

  5. 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.

  6. Results of Small-scale Solid Rocket Combustion Simulator testing at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Goldberg, Benjamin E.; Cook, Jerry

    1993-01-01

    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.

  7. 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.

  8. 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

  9. 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.

  10. 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.

  11. Generic magnetic fusion rocket model

    SciTech Connect

    Santarius, J.F.; Logan, B.G.

    1993-06-01

    A generic magnetic fusion rocket model is developed and used to explore the limits of fusion propulsion systems. Two fusion fuels are examined, D-T and D-(He-3), and the D-(He-3) fuel cycle is found to give a higher specific power in almost all parameter regimes. The key findings are that (1) magnetic fusion should ultimately be able to deliver specific powers of about 10 kW/kg and (2) specific powers of 15 kW/kg could be achieved with only modest extrapolations of present technology. 9 refs.

  12. 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.

  13. Mixing and reaction processes in rocket based combined cycle and conventional rocket engines

    NASA Astrophysics Data System (ADS)

    Lehman, Matthew Kurt

    Raman spectroscopy was used to make species measurements in two rocket engines. An airbreathing rocket, the rocket based combined cycle (RBCC) engine, and a conventional rocket were investigated. A supersonic rocket plume mixing with subsonic coflowing air characterizes the ejector mode of the RBCC engine. The mixing length required for the air and plume to become homogenous is a critical dimension. For the conventional rocket experiments, a gaseous oxygen/gaseous hydrogen single-element shear coaxial injector was used. Three chamber Mach number conditions, 0.1, 0.2 and 0.3, were chosen to assess the effect of Mach number on mixing. The flow within the chamber was entirely subsonic. For the RBCC experiments, vertical Raman line measurements were made at multiple axial locations downstream from the rocket nozzle plane. Species profiles assessed the mixing progress between the supersonic plume and subsonic air. For the conventional rocket, Raman line measurements were made downstream from the injector face. The goal was to evaluate the effect of increased chamber Mach number on injector mixing/reaction. For both engines, quantitative and qualitative information was collected for computational fluid dynamics (CFD development. The RBCC experiments were conducted for three distinct geometries. The primary flow path was a diffuse and afterburner design with a direct-connect air supply. A sea-level static (SLS) version and a thermally choked variant were also tested. The experimental results show that mixing length increases with additional coflow air in the DAB geometry. Operation of variable rocket mixture ratios at identical air flow rates did not significantly affect the mixing length. The thermally choked variant had a longer mixing length compared to the DAB geometry, and the SLS modification had a shorter mixing length due to a reduced air flow. The conventional rocket studies focused on the effect of chamber Mach number on primary injector mixing. Chamber Mach

  14. 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.

  15. 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.

  16. Sounding rockets in Antarctica

    NASA Technical Reports Server (NTRS)

    Alford, G. C.; Cooper, G. W.; Peterson, N. E.

    1982-01-01

    Sounding rockets are versatile tools for scientists studying the atmospheric region which is located above balloon altitudes but below orbital satellite altitudes. Three NASA Nike-Tomahawk sounding rockets were launched from Siple Station in Antarctica in an upper atmosphere physics experiment in the austral summer of 1980-81. The 110 kg payloads were carried to 200 km apogee altitudes in a coordinated project with Arcas rocket payloads and instrumented balloons. This Siple Station Expedition demonstrated the feasibility of launching large, near 1,000 kg, rocket systems from research stations in Antarctica. The remoteness of research stations in Antarctica and the severe environment are major considerations in planning rocket launching expeditions.

  17. 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.

  18. A hybrid rocket engine design for simple low cost sounding rocket use

    NASA Astrophysics Data System (ADS)

    Grubelich, Mark; Rowland, John; Reese, Larry

    1993-06-01

    Preliminary test results on a nitrous oxide/HTPB hybrid rocket engine suitable for powering a small sounding rocket to altitudes of 50-100 K/ft are presented. It is concluded that the advantage of the N2O hybrid engine over conventional solid propellant rocket motors is the ability to obtain long burn times with core burning geometries due to the low regression rate of the fuel. Long burn times make it possible to reduce terminal velocity to minimize air drag losses.

  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. 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.

  1. 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

  2. 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.

  3. 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.

  4. 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.

  5. 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

  6. 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.

  7. 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

  8. 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.

  9. 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.

  10. Supersonic Rocket Thruster Flow Predicted by Numerical Simulation

    NASA Technical Reports Server (NTRS)

    Davoudzadeh, Farhad

    2004-01-01

    Despite efforts in the search for alternative means of energy, combustion still remains the key source. Most propulsion systems primarily use combustion for their needed thrust. Associated with these propulsion systems are the high-velocity hot exhaust gases produced as the byproducts of combustion. These exhaust products often apply uneven high temperature and pressure over the surfaces of the appended structures exposed to them. If the applied pressure and temperature exceed the design criteria of the surfaces of these structures, they will not be able to protect the underlying structures, resulting in the failure of the vehicle mission. An understanding of the flow field associated with hot exhaust jets and the interactions of these jets with the structures in their path is critical not only from the design point of view but for the validation of the materials and manufacturing processes involved in constructing the materials from which the structures in the path of these jets are made. The hot exhaust gases often flow at supersonic speeds, and as a result, various incident and reflected shock features are present. These shock structures induce abrupt changes in the pressure and temperature distribution that need to be considered. In addition, the jet flow creates a gaseous plume that can easily be traced from large distances. To study the flow field associated with the supersonic gases induced by a rocket engine, its interaction with the surrounding surfaces, and its effects on the strength and durability of the materials exposed to it, NASA Glenn Research Center s Combustion Branch teamed with the Ceramics Branch to provide testing and analytical support. The experimental work included the full range of heat flux environments that the rocket engine can produce over a flat specimen. Chamber pressures were varied from 130 to 500 psia and oxidizer-to-fuel ratios (o/f) were varied from 1.3 to 7.5.

  11. 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.

  12. 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.

  13. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

    This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. 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 SpaceTransportation 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.

  14. 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.

  15. 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.

  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. Low thrust chemical rocket technology

    NASA Astrophysics Data System (ADS)

    Schneider, Steven J.

    1992-11-01

    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

  18. 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.

  19. 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.

  20. Life Saving Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    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.

  1. Model Rockets and Microchips.

    ERIC Educational Resources Information Center

    Fitzsimmons, Charles P.

    1986-01-01

    Points out the instructional applications and program possibilities of a unit on model rocketry. Describes the ways that microcomputers can assist in model rocket design and in problem calculations. Provides a descriptive listing of model rocket software for the Apple II microcomputer. (ML)

  2. Postal Rocket Stamps

    NASA Technical Reports Server (NTRS)

    2004-01-01

    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.

  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. Potential Climate and Ozone Impacts From Hybrid Rocket Engine Emissions

    NASA Astrophysics Data System (ADS)

    Ross, M.

    2009-12-01

    Hybrid rocket engines that use N2O as an oxidizer and a solid hydrocarbon (such as rubber) as a fuel are relatively new. Little is known about the composition of such hybrid engine emissions. General principles and visual inspection of hybrid plumes suggest significant soot and possibly NO emissions. Understanding hybrid rocket emissions is important because of the possibility that a fleet of hybrid powered suborbital rockets will be flying on the order of 1000 flights per year by 2020. The annual stratospheric emission for these rockets would be about 10 kilotons, equal to present day solid rocket motor (SRM) emissions. We present a preliminary analysis of the magnitude of (1) the radiative forcing from soot emissions and (2) the ozone depletion from soot and NO emissions associated with such a fleet of suborbital hybrid rockets. Because the details of the composition of hybrid emissions are unknown, it is not clear if the ozone depletion caused by these hybrid rockets would be more or less than the ozone depletion from SRMs. We also consider the climate implications associated with the N2O production and use requirements for hybrid rockets. Finally, we identify the most important data collection and modeling needs that are required to reliably assess the complete range of environmental impacts of a fleet of hybrid rockets.

  5. Liquid-hydrogen rocket engine development at Aerojet, 1944 - 1950

    NASA Technical Reports Server (NTRS)

    Osborn, G. H.; Gordon, R.; Coplen, H. L.; James, G. S.

    1977-01-01

    This program demonstrated the feasibility of virtually all the components in present-day, high-energy, liquid-rocket engines. Transpiration and film-cooled thrust chambers were successfully operated. The first liquid-hydrogen tests of the coaxial injector was conducted and the first pump to successfully produce high pressures in pumping liquid hydrogen was tested. A 1,000-lb-thrust gaseous propellant and a 3,000-lb-thrust liquid-propellant thrust chamber were operated satisfactorily. Also, the first tests were conducted to evaluate the effects of jet overexpansion and separation on performance of rocket thrust chambers with hydrogen-oxygen propellants.

  6. Amateur Gas-Propelled Rocket Engine Development and Advanced Rocket Design

    NASA Astrophysics Data System (ADS)

    Souverein, L. J.; Twigt, D. J.; Engelen, S.

    The paper describes the design and manufacturing of a gaseous propellant rocket engine. It is an undertaking of the authors, performed on project basis with fellow aerospace engineering students under auspices of DARE (Delft Aerospace Rocket Engineering). This paper describes the requirements, the engine development, and the design considerations and calculations as they were performed. Furthermore, the plans for engine tests and the parameters that will have to be measured during those tests are covered. The design process converged to a 1800 N thrust gaseous oxygen/methane (GOX/CH4) engine made of electrolytic copper. GOX/CH4 was selected based on its relatively high specific impulse, its availability and because of its potential as a green propellant. A test engine was produced with a specific impulse of 287 s and a propellant mass flow of 637 g/s. From a point of view of strength, the focus was mainly on robustness rather than light weight. The main aim now is to perform tests with the current engine, based on which the performance can be verified and vital information for future design efforts can be acquired. The ultimate goal is to have an operational rocket and to attempt an amateur altitude record.

  7. Ignition of Hydrogen-Oxygen Rocket Combustor with Chlorine Trifluoride and Triethylaluminum

    NASA Technical Reports Server (NTRS)

    Gregory, John W.; Straight, David M.

    1961-01-01

    Ignition of a nominal-125-pound-thrust cold (2000 R) gaseous-hydrogen - liquid-oxygen rocket combustor with chlorine trifluoride (hypergolic with hydrogen) and triethylaluminum (hypergolic with oxygen) resulted in consistently smooth starting transients for a wide range of combustor operating conditions. The combustor exhaust nozzle discharged into air at ambient conditions. Each starting transient consisted of the following sequence of events: injection of the lead main propellant, injection of the igniter chemical, ignition of these two chemicals, injection of the second main propellant, ignition of the two main propellants, increase in chamber pressure to its terminal value, and cutoff of igniter-chemical flow. Smooth ignition was obtained with an ignition delay of less than 100 milliseconds for the reaction of the lead propellant with the igniter chemical using approximately 0.5 cubic inch (0-038 lb) of chlorine trifluoride or 1.0 cubic inch (0-031 lb) of triethylaluminum. These quantities of igniter chemical were sufficient to ignite a 20-percent-fuel hydrogen-oxygen mixture with a delay time of less than 15 milliseconds. Test results indicated that a simple, light weight chemical ignition system for hydrogen-oxygen rocket engines may be possible.

  8. Improved airbreathing launch vehicle performance with the use of rocket propulsion

    NASA Astrophysics Data System (ADS)

    Kauffman, H. G.; Grandhi, R. V.; Hankey, W. L.; Belcher, P. J.

    1991-04-01

    An efficient performance analysis method is developed to evaluate potential airbreathing/rocket propulsion systems for advanced technology single-stage-to-orbit launch vehicles. Evaluated are tradeoffs between airbreathing, rocket, and concurrent airbreathing/rocket propulsion in maximizing payload delivery to orbit for a given ascent flight trajectory. With the analysis method, several modes of airbreathing/rocket propulsion are compared to a baseline 'airbreather alone' propulsion system in terms of fuel/propellant required to attain orbital velocity. Concurrent airbreathing/rocket propulsion shows a reduction in fuel/propellant consumption over straight airbreather to rocket propulsion transition. The optimal switch point (staging) is identified for the transition from airbreathing to rocket propulsion.

  9. Laser-fusion rocket for interplanetary propulsion

    SciTech Connect

    Hyde, R.A.

    1983-09-27

    A rocket powered by fusion microexplosions is well suited for quick interplanetary travel. Fusion pellets are sequentially injected into a magnetic thrust chamber. There, focused energy from a fusion Driver is used to implode and ignite them. Upon exploding, the plasma debris expands into the surrounding magnetic field and is redirected by it, producing thrust. This paper discusses the desired features and operation of the fusion pellet, its Driver, and magnetic thrust chamber. A rocket design is presented which uses slightly tritium-enriched deuterium as the fusion fuel, a high temperature KrF laser as the Driver, and a thrust chamber consisting of a single superconducting current loop protected from the pellet by a radiation shield. This rocket can be operated with a power-to-mass ratio of 110 W gm/sup -1/, which permits missions ranging from occasional 9 day VIP service to Mars, to routine 1 year, 1500 ton, Plutonian cargo runs.

  10. Gaseous dielectrics V

    SciTech Connect

    Christophorou, L.G.; Bouldin, D.W.

    1987-01-01

    This symposium represents a transdisciplinary and comprehensive approach to the study of gaseous dielectrics. The goal of the symposium was to demonstrate the effective coupling between basic and applied research and modern technology achieved in this area, and to guide future research and development and industrial use of gaseous dielectrics. Separate abstracts were prepared for 85 papers in these proceedings. (DWL)

  11. Impact of rocket exhaust plumes on atmospheric composition and climate ― an overview

    NASA Astrophysics Data System (ADS)

    Voigt, Ch.; Schumann, U.; Graf, K.; Gottschaldt, K.-D.

    2013-03-01

    Rockets are the only direct anthropogenic emission sources into the upper atmosphere. Gaseous rocket emissions include CO, N2, H2, H2O, and CO2, while solid rocket motors (SRM) additionally inject significant amounts of aluminum oxide (Al2O3) particles and gaseous chlorine species into the atmosphere. These emissions strongly perturb local atmospheric trace gas and aerosol distributions. Here, previous aircraft measurements in various rocket exhaust plumes including several large space shuttle launch vehicles are compiled. The observed changes of the lower stratospheric composition in the near field are summarized. The injection of chlorine species and particles into the stratosphere can lead to ozone loss in rocket exhaust plumes. Local observations are compared with global model simulations of the effects of rocket emissions on stratospheric ozone concentrations. Large uncertainties remain concerning individual ozone loss reaction rates and the impact of small-scale plume effects on global chemistry. Further, remote sensing data from satellite indicate that rocket exhaust plumes regionally increase iron and water vapor concentrations in the mesosphere potentially leading to the formation of mesospheric clouds at 80- to 90-kilometer altitude. These satellite observations are summarized and the rocket emission inventory is compared with other natural and anthropogenic sources to the stratosphere such as volcanism, meteoritic material, and aviation.

  12. Indians Repulse British With Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    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.

  13. Another Look at Rocket Thrust

    ERIC Educational Resources Information Center

    Hester, Brooke; Burris, Jennifer

    2012-01-01

    Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force--the thrust--on the rocket. Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of…

  14. 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.

  15. Marshall Team Fires Recreated 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 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. In this photo, the replica is shown firing in the A-frame launch stand in near-flight configuration at MSFC's Test Area 116 during the American Institute of Aeronautics and Astronautics 39th Joint Propulsion Conference on July 23, 2003.

  16. Baking Soda and Vinegar Rockets

    NASA Astrophysics Data System (ADS)

    Claycomb, James R.; Zachary, Christopher; Tran, Quoc

    2009-02-01

    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors1,2 that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the experimentally measured rocket height. Baking soda and vinegar rockets present fewer safety concerns and require a smaller launch area than rapid combustion chemical rockets. Both kits were of nearly identical design, costing ˜20. The rockets required roughly 30 minutes of assembly time consisting of mostly taping the soft plastic fuselage to the Styrofoam nose cone.

  17. A Heated Tube Facility for Rocket Coolant Channel Research

    NASA Technical Reports Server (NTRS)

    Green, James M.; Pease, Gary M.; Meyer, Michael L.

    1995-01-01

    The capabilities of a heated tube facility used for testing rocket engine coolant channels at the NASA Lewis Research Center are presented. The facility uses high current, low voltage power supplies to resistively heat a test section to outer wall temperatures as high as 730 C (1350 F). Liquid or gaseous nitrogen, gaseous helium, or combustible liquids can be used as the test section coolant. The test section is enclosed in a vacuum chamber to minimize heat loss to the surrounding system. Test section geometry, size, and material; coolant properties; and heating levels can be varied to generate heat transfer and coolant performance data bases.

  18. ASTRID rocket flight test

    SciTech Connect

    Whitehead, J.C.; Pittenger, L.C.; Colella, N.J.

    1994-07-01

    On February 4, 1994, we successfully flight tested the ASTRID rocket from Vandenberg Air Force Base. The technology for this rocket originated in the Brilliant Pebbles program and represents a five-year development effort. This rocket demonstrated how our new pumped-propulsion technology-which reduced the total effective engine mass by more than one half and cut the tank mass to one fifth previous requirements-would perform in atmospheric flight. This demonstration paves the way for potential cost-effective uses of the new propulsion system in commercial aerospace vehicles, exploration of the planets, and defense applications.

  19. 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.

  20. Peregrine 100-km Sounding Rocket Project

    NASA Technical Reports Server (NTRS)

    Zilliac, Gregory

    2012-01-01

    The Peregrine Sounding Rocket Program is a joint basic research program of NASA Ames Research Center, NASA Wallops, Stanford University, and the Space Propulsion Group, Inc. (SPG). The goal is to determine the applicability of this technology to a small launch system. The approach is to design, build, and fly a stable, efficient liquefying fuel hybrid rocket vehicle to an altitude of 100 km. The program was kicked off in October of 2006 and has seen considerable progress in the subsequent 18 months. This research group began studying liquifying hybrid rocket fuel technology more than a decade ago. The overall goal of the research was to gain a better understanding of the fundamental physics of the liquid layer entrainment process responsible for the large increase in regression rate observed in these fuels, and to demonstrate the effect of increased regression rate on hybrid rocket motor performance. At the time of this reporting, more than 400 motor tests were conducted with a variety of oxidizers (N2O, GOx, LOx) at ever increasing scales with thrust levels from 5 to over 15,000 pounds (22 N to over 66 kN) in order to move this technology from the laboratory to practical applications. The Peregrine program is the natural next step in this development. A number of small sounding rockets with diameters of 3, 4, and 6 in. (7.6, 10.2, and 15.2 cm) have been flown, but Peregrine at a diameter of 15 in. (38.1 cm) and 14,000-lb (62.3-kN) thrust is by far the largest system ever attempted and will be one of the largest hybrids ever flown. Successful Peregrine flights will set the stage for a wide range of applications of this technology.

  1. Rocket engine numerical simulation

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

  2. Hybrid Rocket Motor Test

    NASA Technical Reports Server (NTRS)

    1994-01-01

    Stennis Space Center conducts a test on a hybrid rocket motor fed by a liquid oxygen turbopump. The test occurred at the E-1 test facility. The test was believed to be the first of its kind in the world.

  3. Antares Rocket Lifts Off!

    NASA Video Gallery

    NASA commercial space partner Orbital Sciences Corp. of Dulles, Va., launched its Cygnus cargo spacecraft aboard its Antares rocket at 10:58 a.m. EDT Wednesday from the Mid-Atlantic Regional Spacep...

  4. Robust Rocket Engine Concept

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.

    1995-01-01

    The potential for a revolutionary step in the durability of reusable rocket engines is made possible by the combination of several emerging technologies. The recent creation and analytical demonstration of life extending (or damage mitigating) control technology enables rapid rocket engine transients with minimum fatigue and creep damage. This technology has been further enhanced by the formulation of very simple but conservative continuum damage models. These new ideas when combined with recent advances in multidisciplinary optimization provide the potential for a large (revolutionary) step in reusable rocket engine durability. This concept has been named the robust rocket engine concept (RREC) and is the basic contribution of this paper. The concept also includes consideration of design innovations to minimize critical point damage.

  5. Rocketing into Adaptive Inquiry.

    ERIC Educational Resources Information Center

    Farenga, Stephen J.; Joyce, Beverly A.; Dowling, Thomas W.

    2002-01-01

    Defines adaptive inquiry and argues for employing this method which allows lessons to be shaped in response to student needs. Illustrates this idea by detailing an activity in which teams of students build rockets. (DDR)

  6. Rocket University at KSC

    NASA Technical Reports Server (NTRS)

    Sullivan, Steven J.

    2014-01-01

    "Rocket University" is an exciting new initiative at Kennedy Space Center led by NASA's Engineering and Technology Directorate. This hands-on experience has been established to develop, refine & maintain targeted flight engineering skills to enable the Agency and KSC strategic goals. Through "RocketU", KSC is developing a nimble, rapid flight engineering life cycle systems knowledge base. Ongoing activities in RocketU develop and test new technologies and potential customer systems through small scale vehicles, build and maintain flight experience through balloon and small-scale rocket missions, and enable a revolving fresh perspective of engineers with hands on expertise back into the large scale NASA programs, providing a more experienced multi-disciplined set of systems engineers. This overview will define the Program, highlight aspects of the training curriculum, and identify recent accomplishments and activities.

  7. Experimental and simulation study of a Gaseous oxygen/Gaseous hydrogen vortex cooling thrust chamber

    NASA Astrophysics Data System (ADS)

    Yu, Nanjia; Zhao, Bo; Li, Gongnan; Wang, Jue

    2016-01-01

    In this paper, RNG k-ε turbulence model and PDF non-premixed combustion model are used to simulate the influence of the diameter of the ring of hydrogen injectors and oxidizer-to-fuel ratio on the specific impulse of the vortex cooling thrust chamber. The simulation results and the experimental tests of a 2000 N Gaseous oxygen/Gaseous hydrogen vortex cooling thrust chamber reveal that the efficiency of the specific impulse improves significantly with increasing of the diameter of the ring of hydrogen injectors. Moreover, the optimum efficiency of the specific impulse is obtained when the oxidizer-to-fuel ratio is near the stoichiometric ratio.

  8. 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. PMID:26078153

  9. Rocket Motor Microphone Investigation

    NASA Technical Reports Server (NTRS)

    Pilkey, Debbie; Herrera, Eric; Gee, Kent L.; Giraud, Jerom H.; Young, Devin J.

    2010-01-01

    At ATK's facility in Utah, large full-scale solid rocket motors are tested. The largest is a five-segment version of the reusable solid rocket motor, which is for use on the Ares I launch vehicle. As a continuous improvement project, ATK and BYU investigated the use of microphones on these static tests, the vibration and temperature to which the instruments are subjected, and in particular the use of vent tubes and the effects these vents have at low frequencies.

  10. Gas Emission Measurements from the RD 180 Rocket Engine

    NASA Technical Reports Server (NTRS)

    Ross, H. R.

    2001-01-01

    The Science Laboratory operated by GB Tech was tasked by the Environmental Office at the NASA Marshall Space Flight Center (MSFC) to collect rocket plume samples and to measure gaseous components and airborne particulates from the hot test firings of the Atlas III/RD 180 test article at MSFC. This data will be used to validate plume prediction codes and to assess environmental air quality issues.

  11. Infrasound Rocket Signatures

    NASA Astrophysics Data System (ADS)

    Olson, J.

    2012-09-01

    This presentation reviews the work performed by our research group at the Geophysical Institute as we have applied the tools of infrasound research to rocket studies. This report represents one aspect of the effort associated with work done for the National Consortium for MASINT Research (NCMR) program operated by the National MASINT Office (NMO) of the Defense Intelligence Agency (DIA). Infrasound, the study of acoustic signals and their propagation in a frequency band below 15 Hz, enables an investigator to collect and diagnose acoustic signals from distant sources. Absorption of acoustic energy in the atmosphere decreases as the frequency is reduced. In the infrasound band signals can propagate hundreds and thousands of kilometers with little degradation. We will present an overview of signatures from rockets ranging from small sounding rockets such as the Black Brandt and Orion series to larger rockets such as Delta 2,4 and Atlas V. Analysis of the ignition transients provides information that can uniquely identify the motor type. After the rocket ascends infrasound signals can be used to characterize the rocket and identify the various events that take place along a trajectory such as staging and maneuvering. We have also collected information on atmospheric shocks and sonic booms from the passage of supersonic vehicles such as the shuttle. This review is intended to show the richness of the unique signal set that occurs in the low-frequency infrasound band.

  12. Rocket Engine Oscillation Diagnostics

    NASA Technical Reports Server (NTRS)

    Nesman, Tom; Turner, James E. (Technical Monitor)

    2002-01-01

    Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

  13. 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.

  14. Guidance and control of an earth to orbit vehicle with optimum transition from airbreathing to concurrent airbreathing/rocket propulsion

    NASA Astrophysics Data System (ADS)

    Kauffman, H. G.; Grandhi, R. V.; Hankey, W. L.; Belcher, P. J.

    1990-07-01

    An efficient performance analysis method (suitable for PC operation) is developed to evaluate potential airbreathing/rocket propulsion systems for advanced technology single-stage-to-orbit (SSTO) launch vehicles. Evaluated are tradeoffs between airbreathing (AB), rocket, and concurrent airbreathing/rocket propulsion in minimizing fuel consumption for a given ascent flight trajectory. Many mission, flight, and vehicle related requirements and constraints are satisfied in the process. With the analysis method, several modes of airbreathing/rocket propulsion are compared to a baseline 'airbreather alone' propulsion system in terms of fuel required to attain orbital velocity. The optimal switch point (staging) is identified for the transition from airbreathing to rocket propulsion.

  15. Hybrid rockets - Combining the best of liquids and solids

    NASA Astrophysics Data System (ADS)

    Cook, Jerry R.; Goldberg, Ben E.; Estey, Paul N.; Wiley, Dan R.

    1992-07-01

    Hybrid rockets employing liquid oxidizer and solid fuel grain answers to cost, safety, reliability, and environmental impact concerns that have become as prominent as performance in recent years. The oxidizer-free grain has limited sensitivity to grain anomalies, such as bond-line separations, which can cause catastrophic failures in solid rocket motors. An account is presently given of the development effort associated with the AMROC commercial hybrid booster and component testing efforts at NASA-Marshall. These hybrid rockets can be fired, terminated, inspected, evaluated, and restarted for additional testing.

  16. Hybrid rockets - Combining the best of liquids and solids

    NASA Technical Reports Server (NTRS)

    Cook, Jerry R.; Goldberg, Ben E.; Estey, Paul N.; Wiley, Dan R.

    1992-01-01

    Hybrid rockets employing liquid oxidizer and solid fuel grain answers to cost, safety, reliability, and environmental impact concerns that have become as prominent as performance in recent years. The oxidizer-free grain has limited sensitivity to grain anomalies, such as bond-line separations, which can cause catastrophic failures in solid rocket motors. An account is presently given of the development effort associated with the AMROC commercial hybrid booster and component testing efforts at NASA-Marshall. These hybrid rockets can be fired, terminated, inspected, evaluated, and restarted for additional testing.

  17. Predicted rocket and shuttle effects on stratospheric ozone

    NASA Technical Reports Server (NTRS)

    Harwood, Robert S.; Karol, Igor L.; Jackman, Charles H.; Qiu, Lian-Xiong; Prather, Michael J.; Pyle, John A.

    1991-01-01

    The major chemical effluents of either solid- or liquid-fueled rockets that can potentially perturb stratospheric ozone include chlorine compounds (HCl), nitrogen compounds (NO(x)), and hydrogen compounds (H2 and H2O). Radicals (Cl, ClO, H, OH, HO2, NO, and NO2) formed directly or indirectly from rocket exhaust can cause the catalytic destruction of ozone. Other exhaust compounds that could presumably lead to ozone destruction either by direct reaction with ozone or by providing a surface for heterogeneous processes include the particulates Al2O3, ice, and soot. These topics are discussed in terms of the possible effects of rocket exhausts on stratospheric ozone.

  18. Some design aspects of laser fusion rocket(II)

    SciTech Connect

    Nakashima, H.; Shoyama, H.; Nagamine, Y.; Fusuki, T.; Kanda, Y.; Yoshimi, N.; Nakao, Y.

    1996-05-01

    A rocket powered by a laser-induced fusion microexplosion is an attractive spacecraft for manned exploration of the solar system because of its potential advantages of high specific impulse and of large specific power. We here discuss some design aspects of a laser fusion rocket. Included are discussion on thrust conversion efficiency, plasma instability in a magnetic nozzle and DT igniter/D{sup 3}He fuel target design. {copyright} {ital 1996 American Institute of Physics.}

  19. 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.

  20. Instrumentation of UALR labscale hybrid rocket motor

    NASA Astrophysics Data System (ADS)

    Wright, Andrew B.; Teague, Warfield; Wright, Ann M.; Wilson, Edmond W.

    2006-05-01

    The Central Arkansas Combustion Group has used a NASA EPSCoR grant to improve the instrumentation and control of its labscale hybrid rocket facility. The research group investigates fundamental aspects of combustion in hybrid rocket motors. This paper describes the new instrumentation, provides examples of measurements taken, and describes novel instrumentation which is in the process of development. A six degree-of-freedom thrust system measures the total work done during a burn to compare the efficiency of fuels and fuel additives. The new system measures the forces and moments in three spatial dimensions. An accurate measure of thrust oscillations will lead to better understanding of the cause and eventual minimization of the oscillations. Plume spectrometers are employed to determine and measure the reaction intermediates and products of combustion at the exhaust. The new control system features an oxygen mass flow controller, which allows the accurate measurement of the oxidant introduced into the motor.

  1. 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

  2. Los Alamos Novel Rocket Design Flight Tested

    ScienceCinema

    Tappan, Bryce

    2015-01-05

    Los Alamos National Laboratory scientists recently flight tested a new rocket design that includes a high-energy fuel and a motor design that also delivers a high degree of safety. Researchers will now work to scale-up the design, as well as explore miniaturization of the system, in order to exploit all potential applications that would require high-energy, high-velocity, and correspondingly high safety margins.

  3. Los Alamos Novel Rocket Design Flight Tested

    SciTech Connect

    Tappan, Bryce

    2014-10-23

    Los Alamos National Laboratory scientists recently flight tested a new rocket design that includes a high-energy fuel and a motor design that also delivers a high degree of safety. Researchers will now work to scale-up the design, as well as explore miniaturization of the system, in order to exploit all potential applications that would require high-energy, high-velocity, and correspondingly high safety margins.

  4. 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. PMID:17339077

  5. 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

  6. Hydrocarbon Rocket Technology Impact Forecasting

    NASA Technical Reports Server (NTRS)

    Stuber, Eric; Prasadh, Nishant; Edwards, Stephen; Mavris, Dimitri N.

    2012-01-01

    Forecasting method is a normative forecasting technique that allows the designer to quantify the effects of adding new technologies on a given design. This method can be used to assess and identify the necessary technological improvements needed to close the gap that exists between the current design and one that satisfies all constraints imposed on the design. The TIF methodology allows for more design knowledge to be brought to the earlier phases of the design process, making use of tools such as Quality Function Deployments, Morphological Matrices, Response Surface Methodology, and Monte Carlo Simulations.2 This increased knowledge allows for more informed decisions to be made earlier in the design process, resulting in shortened design cycle time. This paper will investigate applying the TIF method, which has been widely used in aircraft applications, to the conceptual design of a hydrocarbon rocket engine. In order to reinstate a manned presence in space, the U.S. must develop an affordable and sustainable launch capability. Hydrocarbon-fueled rockets have drawn interest from numerous major government and commercial entities because they offer a low-cost heavy-lift option that would allow for frequent launches1. However, the development of effective new hydrocarbon rockets would likely require new technologies in order to overcome certain design constraints. The use of advanced design methods, such as the TIF method, enables the designer to identify key areas in need of improvement, allowing one to dial in a proposed technology and assess its impact on the system. Through analyses such as this one, a conceptual design for a hydrocarbon-fueled vehicle that meets all imposed requirements can be achieved.

  7. ROCKET PORT CLOSURE

    DOEpatents

    Mattingly, J.T.

    1963-02-12

    This invention provides a simple pressure-actuated closure whereby windowless observation ports are opened to the atmosphere at preselected altitudes. The closure comprises a disk which seals a windowless observation port in rocket hull. An evacuated instrument compartment is affixed to the rocket hull adjacent the inner surface of the disk, while the outer disk surface is exposed to the atmosphere through which the rocket is traveling. The pressure differential between the evacuated instrument compartment and the relatively high pressure external atmosphere forces the disk against the edge of the observation port, thereby effecting a tight seai. The instrument compartment is evacuated to a pressure equal to the atmospheric pressure existing at the altitude at which it is desiretl that the closure should open. When the rocket reaches this preselected altitude, the inwardly directed atmospheric force on the disk is just equaled by the residual air pressure force within the instrument compartment. Consequently, the closure disk falls away and uncovers the open observation port. The separation of the disk from the rocket hull actuates a switch which energizes the mechanism of a detecting instrument disposed within the instrument compartment. (AE C)

  8. The 2003 Goddard Rocket Replica Project: A Reconstruction of the World's First Functional Liquid Rocket System

    NASA Technical Reports Server (NTRS)

    Farr, R. A.; Elam, S. K.; Hicks, G. D.; Sanders, T. M.; London, J. R.; Mayne, A. W.; Christensen, D. L.

    2003-01-01

    As a part of NASA s 2003 Centennial of Flight celebration, engineers and technicians at Marshall Space Flight Center (MSFC), Huntsville, Alabama, in cooperation with the Alabama-Mississippi AIAA Section, have reconstructed historically accurate, functional replicas of Dr. Robert H. Goddard s 1926 first liquid- fuel rocket. The purposes of this project were to clearly understand, recreate, and document the mechanisms and workings of the 1926 rocket for exhibit and educational use, creating a vital resource for researchers studying the evolution of liquid rocketry for years to come. The MSFC team s reverse engineering activity has created detailed engineering-quality drawings and specifications describing the original rocket and how it was built, tested, and operated. Static hot-fire tests, as well as flight demonstrations, have further defined and quantified the actual performance and engineering actual performance and engineering challenges of this major segment in early aerospace history.

  9. Rocket Engine Thrust Chamber Assembly

    NASA Technical Reports Server (NTRS)

    Cornelius, Charles S. (Inventor); Counts, Richard H. (Inventor); Myers, W. Neill (Inventor); Lackey, Jeffrey D. (Inventor); Peters, Warren (Inventor); Shadoan, Michael D. (Inventor); Sparks, David L. (Inventor); Lawrence, Timothy W. (Inventor)

    2001-01-01

    A thrust chamber assembly for liquid fueled rocket engines and the method of making it wherein a two-piece mandrel wrapped with a silica tape saturated with a phenolic resin, the tape extending along the mandrel and covering the combustion chamber portion of the mandrel to the throat portion. The phenolic in the tape is cured and the end of the wrap is machined. The remainder of the mandrel is wrapped with a third silica tape. The resin in the third tape is cured and the assembly is machined. The entire assembly is then wrapped with a tow of graphite fibers wetted with an epoxy resin and, after the epoxy resin is cured, the graphite is machined to final dimensions.

  10. J-2S rocket engine

    NASA Astrophysics Data System (ADS)

    Vilja, J. O.; Briley, G. L.; Murphy, T. H.

    1993-06-01

    The principal design characteristics and features of the J-2S rocket engine, developed as a simpler and more robust version of the J-2 engine, are described. The J-2S is a 265,000-lb vacuum thrust engine that delivers 436 sec vacuum thrust with a nozzle expansion ratio of 40 and operates at a chamber pressure of 1,200 psi. The most unique feature of the J-2S is that it incorporates a main chamber tap-off cycle which eliminates the need for a gas generator. Another simplification for the J-2S is the adoption of a centrifugal fuel turbopump to replace the J-2's axial turbopump. A schematic of the J-2S, engine test results, and performance options are presented.

  11. General view of the Solid Rocket Booster's (SRB) Solid Rocket ...

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

    General view of the Solid Rocket Booster's (SRB) Solid Rocket Motor Segments in the Surge Building of the Rotation Processing and Surge Facility at Kennedy Space Center awaiting transfer to the Vehicle Assembly Building and subsequent mounting and assembly on the Mobile Launch Platform. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  12. Rockets in World War I

    NASA Technical Reports Server (NTRS)

    2004-01-01

    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.

  13. Rocket motor aeroacoustics

    NASA Astrophysics Data System (ADS)

    Hegde, U. G.; Strahle, W. C.

    1983-10-01

    Vibration problems in solid propellant rocket motors are investigated. A class of interior flows modelled to simulate flow conditions inside rocket motor cavities is considered. Turbulence generated pressure fluctuations are shown to consist of two components - acoustic and hydrodynamics. The Bernoulli enthalpy theory of aeroacoustics is employed to extract acoustic pressure spectra from experimentally obtained turbulence data and acoustic impedance values at flow boundaries. The effects of turbulence intensities, sidewall acoustic impedance, axial mass blowing distribution, length to diameter ratio of the cavity and different mass flux on the acoustic pressure level are investigated. Typical pressure levels, under rocket motor conditions, are calculated using the A/B model of propellant response. Estimates of the hydrodynamic component of the pressure fluctuation are provided for the case of fully developed turbulent pipe flow terminated by a choked nozzle.

  14. Vega rocket series of multi-stage amateur's rocket program 1965-1968

    NASA Astrophysics Data System (ADS)

    Kerstein, Aleksander; Krmelj, Miloš

    2003-08-01

    The Astronautical and Rocket Society of Celje (ARSC — Astronavtično in raketno društvo Celje) Slovenia has been involved in experimental programs for students and adults since early in 1962 when the early maned space flight inspired many young people. In the history of ARSC (1962-1999) many project undergone the period 37 years, but one is significant; the PROJECT MULTISTAGE ROCKETS VEGA. The present paper contains chronological and systematical presentation of most rockets, launching and static tests undergone during the period of 1965-1968. VEGA - III - C launching was viewed by some of 500 participants of XVIII International Astronautic Federation Congress, which was held in Belgrade in the former Yugoslavia at that time. Project VEGA, whose main objecture was solid fuel ≫micrograne≪ motor of 100 mm to 160 mm diameter improvements and interconnecting motors in parallel spree and sequentially in stages has been completed with rocket VEGA - IV. This rocket has never been launched and it is still in storage.

  15. Mathematical simulation of hydrogen-oxygen combustion in rocket engines using LOGOS code

    NASA Astrophysics Data System (ADS)

    Betelin, V. B.; Shagaliev, R. M.; Aksenov, S. V.; Belyakov, I. M.; Deryuguin, Yu. N.; Korchazhkin, D. A.; Kozelkov, A. S.; Nikitin, V. F.; Sarazov, A. V.; Zelenskiy, D. K.

    2014-03-01

    Hydrogen-oxygen fuels are very attractive now for rocket engines designers, because this pair is ecology friendly. Computer aided design of new effective and clean hydrogen engines needs mathematical tools for supercomputer modeling of hydrogen-oxygen components mixing and combustion in rocket engines. The paper presents the results of developing, verification and validation of mathematical model making it possible to simulate unsteady processes of ignition and combustion in rocket engines.

  16. Advanced liquid rockets

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    1992-01-01

    A program to substitute iridium coated rhenium for silicide coated niobium in thrust chamber fabrications is reviewed. The life limiting phenomena in each of these material systems is also reviewed. Coating cracking and spalling is not a problem with iridium-coated rhenium as in silicide-coated niobium. Use of the new material system enables an 800 K increase in thruster operating temperature from around 1700 K for niobium to 2500 K for rhenium. Specific impulse iridium-coated rhenium rockets is nominally 20 seconds higher than comparable niobium rockets in the 22 N class and nominally 10 seconds higher in the 440 N class.

  17. 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.

  18. 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.

  19. Baking Soda and Vinegar Rockets

    ERIC Educational Resources Information Center

    Claycomb, James R.; Zachary, Christopher; Tran, Quoc

    2009-01-01

    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the…

  20. Rocket center Peenemuende - Personal memories

    NASA Technical Reports Server (NTRS)

    Dannenberg, Konrad; Stuhlinger, Ernst

    1993-01-01

    A brief history of Peenemuende, the rocket center where Von Braun and his team developed the A-4 (V-2) rocket under German Army auspices, and the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes, is presented. Emphasis is placed on the expansion of operations beginning in 1942.

  1. Analysis of Absorption Spectra of Polycyclic Aromatic Hydrocarbons in Gaseous- and Particle- Phase Emissions from Peat Fuel Combustion Under Controlled Conditions

    NASA Astrophysics Data System (ADS)

    Connolly, J. I.; Samburova, V.; Moosmüller, H.; Khlystov, A.

    2015-12-01

    Biomass and fossil fuel burning processes emit important organic pollutants called polycyclic aromatic hydrocarbons (PAHs) into the atmosphere. Smoldering combustion of peat is one of the largest contributors (up to 70%) of carbonaceous species and, therefore, it may be one of the main sources of these PAHs. PAHs can be detrimental to health, they are known to be potent mutagens and suspected carcinogens. They may also contribute to solar light absorption as the particles absorb in the blue and near ultraviolet (UV) region of the solar spectrum ("brown carbon" species). There is very little knowledge and large ambiguity regarding the contribution of PAHs to optical properties of organic carbon (OC) emitted from smoldering biomass combustion. This study focuses on quantifying and analyzing PAHs emitted from peat smoldering combustion to gain more knowledge on their optical properties. Five peat fuels collected in different regions of the world (Russia, USA) were burned under controlled conditions (e.g., relative humidity, combustion efficiency, fuel-moisture content) at the Desert Research Institute Biomass Burning facility (Reno, NV, USA). Combustion aerosols collected on TIGF filters followed by XAD resin cartridges were extracted and analyzed for gas-phase (semi-volatile) and particle-phase PAHs. Filter and XAD samples were extracted separately with dichloromethane followed by acetone using Accelerated Solvent Extractor (ACE 300, Dionex). To determine absorption properties, absorption spectra of extracts and standard PAHs were recorded between 190 and 900 nm with a UV/VIS spectrophotometer (PerkinElmer, Lambda 650). This poster will discuss the potential contribution of PAHs to brown carbon emitted from peat combustion and give a brief comparison with absorption spectra from biomass burning aerosols.

  2. Rocket-Based Combined Cycle Engine Concept Development

    NASA Technical Reports Server (NTRS)

    Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

    2001-01-01

    The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

  3. 40 CFR 86.137-90 - Dynamometer test run, gaseous and particulate emissions.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 19 2012-07-01 2012-07-01 false Dynamometer test run, gaseous and... New Otto-Cycle Complete Heavy-Duty Vehicles; Test Procedures § 86.137-90 Dynamometer test run, gaseous... first period (505 seconds) is run. (2) Petroleum-fueled and methanol-fueled diesel vehicles. The...

  4. 40 CFR 86.137-90 - Dynamometer test run, gaseous and particulate emissions.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 19 2013-07-01 2013-07-01 false Dynamometer test run, gaseous and... New Otto-Cycle Complete Heavy-Duty Vehicles; Test Procedures § 86.137-90 Dynamometer test run, gaseous... first period (505 seconds) is run. (2) Petroleum-fueled and methanol-fueled diesel vehicles. The...

  5. 40 CFR 86.137-90 - Dynamometer test run, gaseous and particulate emissions.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 18 2011-07-01 2011-07-01 false Dynamometer test run, gaseous and... New Otto-Cycle Complete Heavy-Duty Vehicles; Test Procedures § 86.137-90 Dynamometer test run, gaseous... first period (505 seconds) is run. (2) Petroleum-fueled and methanol-fueled diesel vehicles. The...

  6. Experimental Studies of the Heat Transfer to RBCC Rocket Nozzles for CFD Application to Design Methodologies

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    1999-01-01

    Rocket thrusters for Rocket Based Combined Cycle (RBCC) engines typically operate with hydrogen/oxygen propellants in a very compact space. Packaging considerations lead to designs with either axisymmetric or two-dimensional throat sections. Nozzles tend to be either two- or three-dimensional. Heat transfer characteristics, particularly in the throat, where the peak heat flux occurs, are not well understood. Heat transfer predictions for these small thrusters have been made with one-dimensional analysis such as the Bartz equation or scaling of test data from much larger thrusters. The current work addresses this issue with an experimental program that examines the heat transfer characteristics of a gaseous oxygen (GO2)/gaseous hydrogen (GH2) two-dimensional compact rocket thruster. The experiments involved measuring the axial wall temperature profile in the nozzle region of a water-cooled gaseous oxygen/gaseous hydrogen rocket thruster at a pressure of 3.45 MPa. The wall temperature measurements in the thruster nozzle in concert with Bartz's correlation are utilized in a one-dimensional model to obtain axial profiles of nozzle wall heat flux.

  7. An Ejector Air Intake Design Method for a Novel Rocket-Based Combined-Cycle Rocket Nozzle

    NASA Astrophysics Data System (ADS)

    Waung, Timothy S.

    Rocket-based combined-cycle (RBCC) vehicles have the potential to reduce launch costs through the use of several different air breathing engine cycles, which reduce fuel consumption. The rocket-ejector cycle, in which air is entrained into an ejector section by the rocket exhaust, is used at flight speeds below Mach 2. This thesis develops a design method for an air intake geometry around a novel RBCC rocket nozzle design for the rocket-ejector engine cycle. This design method consists of a geometry creation step in which a three-dimensional intake geometry is generated, and a simple flow analysis step which predicts the air intake mass flow rate. The air intake geometry is created using the rocket nozzle geometry and eight primary input parameters. The input parameters are selected to give the user significant control over the air intake shape. The flow analysis step uses an inviscid panel method and an integral boundary layer method to estimate the air mass flow rate through the intake geometry. Intake mass flow rate is used as a performance metric since it directly affects the amount of thrust a rocket-ejector can produce. The design method results for the air intake operating at several different points along the subsonic portion of the Ariane 4 flight profile are found to under predict mass flow rate by up to 8.6% when compared to three-dimensional computational fluid dynamics simulations for the same air intake.

  8. Multi-dimensional combustor flowfield analyses in gas-gas rocket engine

    NASA Technical Reports Server (NTRS)

    Tsuei, Hsin-Hua; Merkle, Charles L.

    1994-01-01

    The objectives of the present research are to improve design capabilities for low thrust rocket engines through understanding of the detailed mixing and combustions processes. Of particular interest is a small gaseous hydrogen-oxygen thruster which is considered as a coordinated part of an on-going experimental program at NASA LeRC. Detailed computational modeling requires the application of the full three-dimensional Navier Stokes equations, coupled with species diffusion equations. The numerical procedure is performed on both time-marching and time-accurate algorithms and using an LU approximate factorization in time, flux split upwinding differencing in space. The emphasis in this paper is focused on using numerical analysis to understand detailed combustor flowfields, including the shear layer dynamics created between fuel film cooling and the core gas in the vicinity on the nearby combustor wall; the integrity and effectiveness of the coolant film; three-dimensional fuel jets injection/mixing/combustion characteristics; and their impacts on global engine performance.

  9. This Is Rocket Science!

    NASA Astrophysics Data System (ADS)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

  10. Liquid rocket engine turbines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

  11. Hybrid rocket instability

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    The paper provides a brief review of theoretical and experimental studies concerned with hybrid rocket instability. The instabilities discussed include atomization and mixing instabilities, chuffing instabilities, pressure coupled combustion instabilities, and vortex shedding. It is emphasized that the future use of hybrid motor systems as viable design alternatives will depend on a better understanding of hybrid instability.

  12. Hybrid rocket instability

    NASA Astrophysics Data System (ADS)

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

    1993-06-01

    The paper provides a brief review of theoretical and experimental studies concerned with hybrid rocket instability. The instabilities discussed include atomization and mixing instabilities, chuffing instabilities, pressure coupled combustion instabilities, and vortex shedding. It is emphasized that the future use of hybrid motor systems as viable design alternatives will depend on a better understanding of hybrid instability.

  13. Liquid Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

    Comprehensive Liquid Rocket Engine testing is essential to risk reduction for Space Flight. Test capability represents significant national investments in expertise and infrastructure. Historical experience underpins current test capabilities. Test facilities continually seek proactive alignment with national space development goals and objectives including government and commercial sectors.

  14. Water Rocket Workout.

    ERIC Educational Resources Information Center

    Esler, William K.; Sanford, Daniel

    1989-01-01

    Water rockets are used to present Newton's three laws of motion to high school physics students. Described is an outdoor activity which uses four students per group. Provides a launch data sheet to record height, angle of elevation, amount of water used, and launch number. (MVL)

  15. Thiokol Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Graves, S. R.

    2000-01-01

    This paper presents viewgraphs on thiokol solid rocket motors. The topics include: 1) Communications; 2) Military and government intelligence; 3) Positioning satellites; 4) Remote sensing; 5) Space burial; 6) Science; 7) Space manufacturing; 8) Advertising; 9) Space rescue space debris management; 10) Space tourism; 11) Space settlements; 12) Hazardous waste disposal; 13) Extraterrestrial resources; 14) Fast package delivery; and 15) Space utilities.

  16. This "Is" Rocket Science!

    ERIC Educational Resources Information Center

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

  17. Liquid rocket valve components

    NASA Technical Reports Server (NTRS)

    1973-01-01

    A monograph on valves for use with liquid rocket propellant engines is presented. The configurations of the various types of valves are described and illustrated. Design criteria and recommended practices for the various valves are explained. Tables of data are included to show the chief features of valve components in use on operational vehicles.

  18. Liquid rocket valve assemblies

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The design and operating characteristics of valve assemblies used in liquid propellant rocket engines are discussed. The subjects considered are as follows: (1) valve selection parameters, (2) major design aspects, (3) design integration of valve subassemblies, and (4) assembly of components and functional tests. Information is provided on engine, stage, and spacecraft checkout procedures.

  19. The Relativistic Rocket

    ERIC Educational Resources Information Center

    Antippa, Adel F.

    2009-01-01

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful…

  20. Solid rocket motors

    NASA Technical Reports Server (NTRS)

    Carpenter, Ronn L.

    1993-01-01

    Structural requirements, materials and, especially, processing are critical issues that will pace the introduction of new types of solid rocket motors. Designers must recognize and understand the drivers associated with each of the following considerations: (1) cost; (2) energy density; (3) long term storage with use on demand; (4) reliability; (5) safety of processing and handling; (6) operability; and (7) environmental acceptance.

  1. Dr. Goddard Transports Rocket

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Dr. Robert H. Goddard tows his rocket to the launching tower behind a Model A Ford truck, 15 miles northwest of Roswell, New Mexico. 1930- 1932. Dr. Goddard has been recognized as the 'Father of American Rocketry' and as one of three pioneers in the theoretical exploration of space. Robert Hutchings Goddard was born in Worcester, Massachusetts, on October 15, 1882. He was a theoretical scientist as well as a practical engineer. His dream was the conquest of the upper atmosphere and ultimately space through the use of rocket propulsion. Dr. Goddard, who died in 1945, was probably as responsible for the dawning of the Space Age as the Wright Brothers were for the begining of the Air Age. Yet his work attracted little serious attention during his lifetime. When the United States began to prepare for the conquest of space in the 1950's, American rocket scientists began to recognize the debt owed to the New England professor. They discovered that it was virtually impossible to construct a rocket or launch a satellite without acknowledging the work of Dr. Goddard. This great legacy was covered by more than 200 patents, many of which were issued after his death.

  2. 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.

  3. Dynamics of the flammable plumes resulting from the convective dispersion of a fixed mass of the buoyant gaseous fuel, methane, into air.

    PubMed

    Fardisi, S; Karim, Ghazi A

    2009-08-15

    The dynamics of the dispersion of a fixed mass of the buoyant fuel, methane, when exposed with a negligible pressure difference to overlaying air within vertical cylindrical enclosures open to the atmosphere is investigated. Features of the formation and dispersion of flammable mixtures created by the gas dissipation were examined using a 3D CFD model. For the cases considered, the lean-flammable mixture boundary appears to travel mainly at a near constant rate while the rich limit front shows a more chaotic behaviour. The corresponding simulation using an axis-symmetrical 2D model tended to under-predict the dynamics of the lean and rich boundaries, for the cases considered. PMID:19237243

  4. Particle bed reactor nuclear rocket concept

    NASA Technical Reports Server (NTRS)

    Ludewig, Hans

    1991-01-01

    The particle bed reactor nuclear rocket concept consists of fuel particles (in this case (U,Zr)C with an outer coat of zirconium carbide). These particles are packed in an annular bed surrounded by two frits (porous tubes) forming a fuel element; the outer one being a cold frit, the inner one being a hot frit. The fuel element are cooled by hydrogen passing in through the moderator. These elements are assembled in a reactor assembly in a hexagonal pattern. The reactor can be either reflected or not, depending on the design, and either 19 or 37 elements, are used. Propellant enters in the top, passes through the moderator fuel element and out through the nozzle. Beryllium used for the moderator in this particular design to withstand the high radiation exposure implied by the long run times.

  5. 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.

  6. Dual Expander Cycle Rocket Engine with an Intermediate, Closed-cycle Heat Exchanger

    NASA Technical Reports Server (NTRS)

    Greene, William D. (Inventor)

    2008-01-01

    A dual expander cycle (DEC) rocket engine with an intermediate closed-cycle heat exchanger is provided. A conventional DEC rocket engine has a closed-cycle heat exchanger thermally coupled thereto. The heat exchanger utilizes heat extracted from the engine's fuel circuit to drive the engine's oxidizer turbomachinery.

  7. 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.

  8. 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.

  9. A performance comparison of two small rocket nozzles

    NASA Technical Reports Server (NTRS)

    Arrington, Lynn A.; Reed, Brian D.; Rivera, Angel, Jr.

    1996-01-01

    An experimental study was conducted on two small rockets (110 N thrust class) to directly compare a standard conical nozzle with a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds numbers of greater than 1 x 10(exp 5), bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below 1 x 10(exp 5), however, test results have been ambiguous. An experimental program was conducted to test two small nozzles at two different fuel film cooling percentages and three different chamber pressures. Test results showed that for the throat Reynolds number range from 2 x 10(exp 4) to 4 x 10(exp 4), the bell nozzle outperformed the conical nozzle. Thrust coefficients for the bell nozzle were approximately 4 to 12 percent higher than those obtained with the conical nozzle. As expected, testing showed that lowering the fuel film cooling increased performance for both nozzle types.

  10. Laser rocket system analysis

    NASA Technical Reports Server (NTRS)

    Jones, W. S.; Forsyth, J. B.; Skratt, J. P.

    1979-01-01

    The laser rocket systems investigated in this study were for orbital transportation using space-based, ground-based and airborne laser transmitters. The propulsion unit of these systems utilizes a continuous wave (CW) laser beam focused into a thrust chamber which initiates a plasma in the hydrogen propellant, thus heating the propellant and providing thrust through a suitably designed nozzle and expansion skirt. The specific impulse is limited only by the ability to adequately cool the thruster and the amount of laser energy entering the engine. The results of the study showed that, with advanced technology, laser rocket systems with either a space- or ground-based laser transmitter could reduce the national budget allocated to space transportation by 10 to 345 billion dollars over a 10-year life cycle when compared to advanced chemical propulsion systems (LO2-LH2) of equal capability. The variation in savings depends upon the projected mission model.

  11. Solid rocket combustion simulator technology using the hybrid rocket for simulation

    NASA Technical Reports Server (NTRS)

    Ramohalli, Kumar

    1994-01-01

    The hybrid rocket is reexamined in light of several important unanswered questions regarding its performance. The well-known heat transfer limited burning rate equation is quoted, and its limitations are pointed out. Several inconsistencies in the burning rate determination through fuel depolymerization are explicitly discussed. The resolution appears to be through the postulate of (surface) oxidative degradation of the fuel. Experiments are initiated to study the fuel degradation in mixtures of nitrogen/oxygen in the 99.9 percent/0.1 percent to 98 percent/2 percent range. The overall hybrid combustion behavior is studied in a 2 in-diameter rocket motor, where a PMMA tube is used as the fuel. The results include detailed, real-time infrared video images of the combustion zone. Space- and time-averaged images give a broad indication of the temperature reached in the gases. A brief outline is shown of future work, which will specifically concentrate on the exploration of the role of the oxidizer transport to the fuel surface, and the role of the unburned fuel that is reported to escape below the classical time-averaged boundary layer flame.

  12. Purge Monitoring Technology for Gaseous Helium (GHe) Conservation

    NASA Technical Reports Server (NTRS)

    Dickey, Jonathan; Lansaw, John

    2010-01-01

    John C. Stennis Space Center provides rocket engine propulsion testing for the NASA space programs. Since the development of the Space Shuttle, every Space Shuttle Main Engine (SSME) has gone through acceptance testing before going to Kennedy Space Center for integration into the Space Shuttle. The SSME is a large cryogenic rocket engine that used Liquid Oxygen (LO2) and Liquid Hydrogen (LH2) as propellants. Due to the extremely cold cryogenic conditions of this environment, an inert gas, helium, is used as a purge for the engine and propellant lines since it can be used without freezing in the cryogenic environment. As NASA moves forward with the development of the new ARES V launch system, the main engines as well as the upper stage engine will use cryogenic propellants and will require gaseous helium during the development testing of each of these engines. The main engine for the ARES V will be similar in size to the SSME.

  13. Small rocket tornado probe

    SciTech Connect

    Colgate, S.A.

    1982-01-01

    A (less than 1 lb.) paper rock tornado probe was developed and deployed in an attempt to measure the pressure, temperature, ionization, and electric field variations along a trajectory penetrating a tornado funnel. The requirements of weight and materials were set by federal regulations and a one-meter resolution at a penetration velocity of close to Mach 1 was desired. These requirements were achieved by telemetering a strain gage transducer for pressure, micro size thermister and electric field, and ionization sensors via a pulse time telemetry to a receiver on board an aircraft that digitizes a signal and presents it to a Z80 microcomputer for recording on mini-floppy disk. Recording rate was 2 ms for 8 channels of information that also includes telemetry rf field strength, magnetic field for orientation on the rocket, zero reference voltage for the sensor op amps as well as the previously mentioned items also. The absolute pressure was recorded. Tactically, over 120 h were flown in a Cessna 210 in April and May 1981, and one tornado was encountered. Four rockets were fired at this tornado, missed, and there were many equipment problems. The equipment needs to be hardened and engineered to a significant degree, but it is believed that the feasibility of the probe, tactics, and launch platform for future tornado work has been proven. The logistics of thunderstorm chasing from a remote base in New Mexico is a major difficulty and reliability of the equipment another. Over 50 dummy rockets have been fired to prove trajectories, stability, and photographic capability. Over 25 electronically equipped rockets have been fired to prove sensors transmission, breakaway connections, etc. The pressure recovery factor was calibrated in the Air Force Academy blow-down tunnel. There is a need for more refined engineering and more logistic support.

  14. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2004-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

  15. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2003-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components and with appropriate adjustment of curing and other additives functionally-required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g. powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf life characteristics.

  16. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2008-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

  17. Excitation by rockets

    NASA Technical Reports Server (NTRS)

    Tammadge, C. E.

    1975-01-01

    Standard methods of excitation are not always practical when a single mode of known frequency requires investigation. This form of investigation is often required on a modified aircraft. A new method of excitation was developed and proved in flight, which consists of firing small rocket charges attached to the aircraft structure. Damping values at gradually increasing airspeeds are obtained, as in Stick Jerk tests, and flutter speeds predicted.

  18. Solid propellant rocket motor

    NASA Technical Reports Server (NTRS)

    Dowler, W. L.; Shafer, J. I.; Behm, J. W.; Strand, L. D. (Inventor)

    1973-01-01

    The characteristics of a solid propellant rocket engine with a controlled rate of thrust buildup to a desired thrust level are discussed. The engine uses a regressive burning controlled flow solid propellant igniter and a progressive burning main solid propellant charge. The igniter is capable of operating in a vacuum and sustains the burning of the propellant below its normal combustion limit until the burning propellant surface and combustion chamber pressure have increased sufficiently to provide a stable chamber pressure.

  19. Two-Dimensional Motions of Rockets

    ERIC Educational Resources Information Center

    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…

  20. Developing a gas rocket performance prediction technique

    NASA Technical Reports Server (NTRS)

    Morgenthaler, J. H.; Moon, L. F.; Stepien, W. R.

    1974-01-01

    A simple, semi-empirical performance correlation/prediction technique applicable to gaseous and liquid propellant rocket engines is presented. Excellent correlations were attained for over 100 test firings by adjusting the computation of the gaseous mixing of an unreactive, coaxial jet using a correlation factor, F, which resulted in prediction of the experimental combustion efficiency for each firing. Static pressure, mean velocity and turbulence intensity in the developing region of non-reactive coaxial jets, typical of those of coaxial injector elements were determined. Detailed profiles were obtained at twelve axial locations (extending from the nozzle exit for a distance of five diameters) downstream from a single element of the Bell Aerospace H2/O2 19-element coaxial injector. These data are compared with analytical predictions made using both eddy viscosity and turbulence kinetic energy mixing models and available computer codes. Comparisons were disappointing, demonstrating the necessity of developing improved turbulence models and computational techniques before detailed predictions of practical coaxial free jet flows are attempted.

  1. Isothermal Gaseous Detonation Model

    NASA Astrophysics Data System (ADS)

    Prokhorov, E. S.

    2015-05-01

    We propose an isothermal gaseous detonation model taking into account the initial pressure of the explosive mixture that permits describing in a simplified form both the self-sustaining and the supercompressed and undercompressed detonation regimes. The exactness of this model has been estimated on the basis of a comparative analysis with the results of equilibrium calculations of the gas-dynamic parameters at the front of detonation waves.

  2. 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)

  3. 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.

  4. Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

    NASA Astrophysics Data System (ADS)

    Elliott, T. S.; Majdalani, J.

    2014-11-01

    Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion.

  5. NEW EMPLOYEES ON THE JOB - DONALD E HEGBERG OF THE NUCLEAR REACTOR DIVISION DISCUSSES NUCLEAR ROCKET

    NASA Technical Reports Server (NTRS)

    1963-01-01

    NEW EMPLOYEES ON THE JOB - DONALD E HEGBERG OF THE NUCLEAR REACTOR DIVISION DISCUSSES NUCLEAR ROCKET FUEL ELEMENT EXPERIMENT WITH CHARLES L YOUNGER - THE DISCUSSION IS PREPATORY TO CONDUCTING THE EXPERIMENT AT THE PLUM BROOK STATION REACTOR FACILITY

  6. 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

  7. Advancing the State-of-the-Practice for Liquid Rocket Engine Injector Design

    NASA Technical Reports Server (NTRS)

    Tucker, P. K.; Kenny, R. J.; Richardson, B. R.; Anderso, W. E.; Austin, B. J.; Schumaker, S. A.; Muss, J. A.

    2015-01-01

    Current shortcomings in both the overall injector design process and its underlying combustion stability assessment methodology are rooted in the use of empirically based or low fidelity representations of complex physical phenomena and geometry details that have first order effects on performance, thermal environments and combustion stability. The result is a design and analysis capability that is often inadequate to reliably arrive at a suitable injector design in an efficient manner. Specifically, combustion instability has been particularly difficult to predict and mitigate. Large hydrocarbon-fueled booster engines have been especially problematic in this regard. Where combustion instability has been a problem, costly and time-consuming redesign efforts have often been an unfortunate consequence. This paper presents an overview of a recently completed effort at NASA Marshall Space Flight Center to advance the state-of-the-practice for liquid rocket engine injector design. Multiple perturbations of a gas-centered swirl coaxial (GCSC) element that burned gaseous oxygen and RP-1 were designed, assessed for combustion stability, and tested. Three designs, one stable, one marginally unstable and one unstable, were used to demonstrate both an enhanced overall injector design process and an improved combustion stability assessment process. High-fidelity results from state-of-the-art computational fluid dynamics CFD simulations were used to substantially augment and improve the injector design methodology. The CFD results were used to inform and guide the overall injector design process. They were also used to upgrade selected empirical or low-dimensional quantities in the ROCket Combustor Interactive Design (ROCCID) stability assessment tool. Hot fire single element injector testing was used to verify both the overall injector designs and the stability assessments. Testing was conducted at the Air Force Research Laboratory and at Purdue University. Companion papers

  8. Delta II rocket prepared for launch of Deep Space 1

    NASA Technical Reports Server (NTRS)

    1998-01-01

    A Boeing Delta 7326 rocket with two solid rocket boosters attached sits on Launch Pad 17A, Cape Canaveral Air Station. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches. Delta's origins go back to the Thor intermediate-range ballistic missile, which was developed in the mid-1950s for the U.S. Air Force. The Thor -- a single-stage, liquid-fueled rocket -- later was modified to become the Delta launch vehicle. Delta IIs are manufactured in Huntington Beach, Calif. Rocketdyne, a division of The Boeing Company, builds Delta II's main engine in Canoga Park, Calif. Final assembly takes place at the Boeing facility in Pueblo, Colo. The Delta 7236, which has three solid rocket boosters and a Star 37 upper stage, will launch Deep Space 1, the first flight in NASA's New Millennium Program. It is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include an ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

  9. Delta II rocket prepared for launch of Deep Space 1

    NASA Technical Reports Server (NTRS)

    1998-01-01

    - A solid rocket booster is maneuvered into place for installation on the Boeing Delta 7326 rocket that will launch Deep Space 1 at Launch Pad 17A, Cape Canaveral Air Station. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches. Delta's origins go back to the Thor intermediate-range ballistic missile, which was developed in the mid-1950s for the U.S. Air Force. The Thor -- a single-stage, liquid-fueled rocket -- later was modified to become the Delta launch vehicle. The Delta 7236 has three solid rocket boosters and a Star 37 upper stage. Delta IIs are manufactured in Huntington Beach, Calif. Rocketdyne, a division of The Boeing Company, builds Delta II's main engine in Canoga Park, Calif. Final assembly takes place at the Boeing facility in Pueblo, Colo. Deep Space 1, the first flight in NASA's New Millennium Program, is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include an ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

  10. Delta II rocket prepared for launch of Deep Space 1

    NASA Technical Reports Server (NTRS)

    1998-01-01

    (Left) A solid rocket booster is lifted for installation onto the Boeing Delta 7326 rocket that will launch Deep Space 1 at Launch Pad 17A, Cape Canaveral Air Station. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches. Delta's origins go back to the Thor intermediate-range ballistic missile, which was developed in the mid-1950s for the U.S. Air Force. The Thor -- a single-stage, liquid-fueled rocket -- later was modified to become the Delta launch vehicle. The Delta 7236 has three solid rocket boosters and a Star 37 upper stage. Delta IIs are manufactured in Huntington Beach, Calif. Rocketdyne, a division of The Boeing Company, builds Delta II's main engine in Canoga Park, Calif. Final assembly takes place at the Boeing facility in Pueblo, Colo. Deep Space 1, the first flight in NASA's New Millennium Program, is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include an ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

  11. Delta II rocket prepared for launch of Deep Space 1

    NASA Technical Reports Server (NTRS)

    1998-01-01

    A solid rocket booster (left) is raised for installation onto the Boeing Delta 7326 rocket that will launch Deep Space 1 at Launch Pad 17A, Cape Canaveral Air Station. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches. Delta's origins go back to the Thor intermediate-range ballistic missile, which was developed in the mid-1950s for the U.S. Air Force. The Thor -- a single-stage, liquid-fueled rocket -- later was modified to become the Delta launch vehicle. The Delta 7236 has three solid rocket boosters and a Star 37 upper stage. Delta IIs are manufactured in Huntington Beach, Calif. Rocketdyne, a division of The Boeing Company, builds Delta II's main engine in Canoga Park, Calif. Final assembly takes place at the Boeing facility in Pueblo, Colo. Deep Space 1, the first flight in NASA's New Millennium Program, is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include an ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

  12. Experimental study of combustion in hydrogen peroxide hybrid rockets

    NASA Astrophysics Data System (ADS)

    Wernimont, Eric John

    Combustion behavior in a hydrogen peroxide oxidized hybrid rocket motor is investigated with a series of experiments. Hybrid chemical rocket propulsion is presently of interest due to reduced system complexity compared to classical chemical propulsion systems. Reduced system complexity, by use of a storable oxidizer and a hybrid configuration, is expected to reduce propulsive costs. The fuel in this study is polyethylene which has the potential of continuous manufacture leading to further reduced system costs. The study investigated parameters of interest for nominal design of a full scale hydrogen peroxide oxidized hybrid rocket. Amongst these parameters is the influence of chamber pressure, mass flux, fuel molecular weight and fuel density on fuel regression rate. Effects of chamber pressure and aft combustion length on combustion efficiency and non-acoustic combustion oscillations are also examined. The fuel regression behavior is found to be strongly influenced by both chamber pressure and mass flux. Combustion efficiencies in the upper 90% range are attained by simple changes to the aft combustion chamber length as well as increased combustion pressure. Fuel burning surface is found to be influenced by the density of the polyethylene polymer as well as molecular weight. The combustion is observed to be exceptionally smooth (oscillations less than 5% zero-to-peak of mean) in all motors tested in this program. Tests using both a single port fuel gain and a novel radial flow hybrid are also performed.

  13. Liquid Rocket Engine Testing Overview

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

    Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

  14. Partial least squares analysis of rocket propulsion fuel data using diaphragm valve-based comprehensive two-dimensional gas chromatography coupled with flame ionization detection.

    PubMed

    Freye, Chris E; Fitz, Brian D; Billingsley, Matthew C; Synovec, Robert E

    2016-06-01

    The chemical composition and several physical properties of RP-1 fuels were studied using comprehensive two-dimensional (2D) gas chromatography (GC×GC) coupled with flame ionization detection (FID). A "reversed column" GC×GC configuration was implemented with a RTX-wax column on the first dimension ((1)D), and a RTX-1 as the second dimension ((2)D). Modulation was achieved using a high temperature diaphragm valve mounted directly in the oven. Using leave-one-out cross-validation (LOOCV), the summed GC×GC-FID signal of three compound-class selective 2D regions (alkanes, cycloalkanes, and aromatics) was regressed against previously measured ASTM derived values for these compound classes, yielding root mean square errors of cross validation (RMSECV) of 0.855, 0.734, and 0.530mass%, respectively. For comparison, using partial least squares (PLS) analysis with LOOCV, the GC×GC-FID signal of the entire 2D separations was regressed against the same ASTM values, yielding a linear trend for the three compound classes (alkanes, cycloalkanes, and aromatics), yielding RMSECV values of 1.52, 2.76, and 0.945 mass%, respectively. Additionally, a more detailed PLS analysis was undertaken of the compounds classes (n-alkanes, iso-alkanes, mono-, di-, and tri-cycloalkanes, and aromatics), and of physical properties previously determined by ASTM methods (such as net heat of combustion, hydrogen content, density, kinematic viscosity, sustained boiling temperature and vapor rise temperature). Results from these PLS studies using the relatively simple to use and inexpensive GC×GC-FID instrumental platform are compared to previously reported results using the GC×GC-TOFMS instrumental platform. PMID:27130110

  15. Rocket engine heat transfer and material technology for commercial applications

    NASA Technical Reports Server (NTRS)

    Hiltabiddle, J.; Campbell, J.

    1974-01-01

    Liquid fueled rocket engine combustion, heat transfer, and material technology have been utilized in the design and development of compact combustion and heat exchange equipment intended for application in the commercial field. An initial application of the concepts to the design of a compact steam generator to be utilized by electrical utilities for the production of peaking power is described.

  16. U.S./CIS eye joint nuclear rocket venture

    NASA Technical Reports Server (NTRS)

    Clark, John S.; Mcilwain, Melvin C.; Smetanikov, Vladimir; D'Yakov, Evgenij K.; Pavshuk, Vladimir A.

    1993-01-01

    An account is given of the significance for U.S. spacecraft development of a nuclear thermal rocket (NTR) reactor concept that has been developed in the (formerly Soviet) Commonwealth of Independent States (CIS). The CIS NTR reactor employs a hydrogen-cooled zirconium hydride moderator and ternary carbide fuels; the comparatively cool operating temperatures associated with this design promise overall robustness.

  17. ISS Update: VASIMR Plasma Rocket

    NASA Video Gallery

    NASA Public Affairs Officer Dan Huot interviews Ken Bollweg, VASIMR Project Manager, about VASIMR (Variable Specific Impulse Magnetoplasma Rocket), recent testing progress and future applications. ...

  18. Electric rockets get a boost

    SciTech Connect

    Ashley, S.

    1995-12-01

    This article reports that xenon-ion thrusters are expected to replace conventional chemical rockets in many nonlaunch propulsion tasks, such as controlling satellite orbits and sending space probes on long exploratory missions. The space age dawned some four decades ago with the arrival of powerful chemical rockets that could propel vehicles fast enough to escape the grasp of earth`s gravity. Today, chemical rocket engines still provide the only means to boost payloads into orbit and beyond. The less glamorous but equally important job of moving vessels around in space, however, may soon be assumed by a fundamentally different rocket engine technology that has been long in development--electric propulsion.

  19. Walter Thiel—Short life of a rocket scientist

    NASA Astrophysics Data System (ADS)

    Thiel, Karen; Przybilski, Olaf

    2013-10-01

    In 2012 we celebrate the 70th anniversary of the first successful rocket launch that reached a height of 84.5 km and had a speed of 4.824 km/h (5x sonic speed). This rocket flew 190 km to the target location. One of the masterminds of this launch was Walter Thiel, a German chemist and rocket engineer. Thiel was highly talented, during his education from primary school until diploma exams he always received a grade of A in his exams. He was called "the student with the 7 A grades". In 1934 Thiel became Dr.-Ing. (chem.), with the highest possible honor (summa cum laude), when he was only 24 years old. He started to work for the rocket development department at Humboldt University, Berlin. Walter Dornberger asked him to leave the university research department and become head of rocket propulsion development in his team in Kummersdorf, near Berlin. Thiel's groundbreaking ideas for the rocket engine would lead to a significant reduction in material, weight and work processes, as well as a shortening in the length of the engine itself. Thiel and his team also defined the fuel itself and the best ratio of mixture between ethanol and liquid oxygen for the engine. In 1940 the propulsion team moved from Kummersdorf to Peenemünde after the launch sites were completed there. Thiel became deputy of Wernher von Braun at the R&D units. One of Thiel's team members was Konrad Dannenberg, who later became famous in the development of the Saturn program. On the night from August 17 to August 18, 1943, Thiel and his family (wife and two children) were killed during a Royal Air Force bombing raid (Operation Hydra). The Moon crater "Thiel" on the far side of the Moon is named after Walter Thiel. The research results of Walter Thiel had a strong impact on the United States' rocket program as well as the Russian rocket development program.

  20. Thermohydraulic modeling of the nuclear thermal rocket: The KLAXON code

    SciTech Connect

    Hall, M.L.; Rider, W.J.; Cappiello, M.W. )

    1992-01-01

    Nuclear thermal rockets (NTRs) have been proposed as a means of propulsion for the Space Exploration Initiative (SEI, the manned mission to Mars). The NTR derives its thrust from the expulsion of hot supersonic hydrogen gas. A large tank on the rocket stores hydrogen in liquid or slush form, which is pumped by a turbopump through a nuclear reactor to provide the necessary heat. The path that the hydrogen takes is most circuitous, making several passes through the reactor and the nozzle itself (to provide cooling), as well as two passes through the turbopump (to transfer momentum). The proposed fuel elements for the reactor have two different configurations: solid prismatic fuel and particle-bed fuel. There are different design concerns for the two types of fuel, but there are also many fluid flow aspects that they share. The KLAXON code was used to model a generic NTR design from the inlet of the reactor core to the exit from the nozzle.

  1. Computational simulation of liquid rocket injector anomalies

    NASA Technical Reports Server (NTRS)

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.; Davidian, K.

    1986-01-01

    A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

  2. Nuclear Thermal Rocket Simulation in NPSS

    NASA Technical Reports Server (NTRS)

    Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas L.

    2013-01-01

    Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

  3. Nuclear Thermal Rocket Simulation in NPSS

    NASA Technical Reports Server (NTRS)

    Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

    2013-01-01

    Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

  4. If Only Newton Had a Rocket.

    ERIC Educational Resources Information Center

    Hammock, Frank M.

    1988-01-01

    Shows how model rocketry can be included in physics curricula. Describes rocket construction, a rocket guide sheet, calculations and launch teams. Discusses the relationships of basic mechanics with rockets. (CW)

  5. Micro-Rockets for the Classroom.

    ERIC Educational Resources Information Center

    Huebner, Jay S.; Fletcher, Alice S.; Cato, Julia A.; Barrett, Jennifer A.

    1999-01-01

    Compares micro-rockets to commercial models and water rockets. Finds that micro-rockets are more advantageous because they are constructed with inexpensive and readily available materials and can be safely launched indoors. (CCM)

  6. Ignition and Flame Stabilization of a Strut-Jet RBCC Combustor with Small Rocket Exhaust

    PubMed Central

    2014-01-01

    A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

  7. Ignition and flame stabilization of a strut-jet RBCC combustor with small rocket exhaust.

    PubMed

    Hu, Jichao; Chang, Juntao; Bao, Wen

    2014-01-01

    A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

  8. Air breathing engine/rocket trajectory optimization

    NASA Technical Reports Server (NTRS)

    Smith, V. K., III

    1979-01-01

    This research has focused on improving the mathematical models of the air-breathing propulsion systems, which can be mated with the rocket engine model and incorporated in trajectory optimization codes. Improved engine simulations provided accurate representation of the complex cycles proposed for advanced launch vehicles, thereby increasing the confidence in propellant use and payload calculations. The versatile QNEP (Quick Navy Engine Program) was modified to allow treatment of advanced turboaccelerator cycles using hydrogen or hydrocarbon fuels and operating in the vehicle flow field.

  9. Creation of a market for small rocket experiments through CAMUI hybrid rocket

    NASA Astrophysics Data System (ADS)

    Nagata, Harunori; Watanabe, Mikio; Ito, Mitsunori; Maeda, Takenori; Uematsu, Tsutomu; Totani, Tsuyoshi; Kudo, Isao

    2005-08-01

    By introducing various innovative ideas, the difficult-to-develop small hybrid-type rocket is successfully developed. The main purpose is to drastically reduce the cost of rocket experiments and thus attract potential users such as metrological and microgravity researchers. A key idea is a new fuel grain design to accelerate the gasification rate of solid fuel. The new fuel grain design, designated as CAMUI as an abbreviation of "Cascaded Multistage Impinging-jet", is that the gas flow repeatedly collides with the solid fuel surface to accelerate the heat transfer to the fuel. To install a regenerative cooling system using cryogenic liquid oxygen as coolant in a small launcher, the authors devised a valveless supply system (with no valves in the liquid oxygen flow line). Four serial successful launch verification tests by 10 kg vehicle equipped with a 50 kgf thrust CAMUI motor have shown the feasibility of the motor system. The meteorological observation model of 400 kgf class motor is under development and the development of microgravity experiment class of 1.5 to 2 tonf motor will follow subsequently. The authors plan to complete the developoment of the 400 kgf class motor for meterological observation model by the end of FY2005.

  10. Development of CAMUI hybrid rocket to create a market for small rocket experiments

    NASA Astrophysics Data System (ADS)

    Nagata, Harunori; Ito, Mitsunori; Maeda, Takenori; Watanabe, Mikio; Uematsu, Tsutomu; Totani, Tsuyoshi; Kudo, Isao

    2006-07-01

    By introducing various innovative ideas, the difficult-to-develop small hybrid-type rocket is successfully developed. The main purpose is to drastically reduce the cost of rocket experiments and thus, attract potential users such as metrological and microgravity researchers. A key idea is a new fuel grain design to accelerate the gasification rate of solid fuel. The new fuel grain design, designated as CAMUI as an abbreviation of "cascaded multistage impinging-jet", is that the gas flow repeatedly collides with the solid fuel surface to accelerate the heat transfer to the fuel. To install a regenerative cooling system using cryogenic liquid oxygen as coolant in a small launcher, the authors devised a valveless supply system (with no valves in the liquid oxygen flow line). Four serial successful launch verification tests by 10 kg vehicle equipped with a 50 kgf thrust CAMUI motor have shown the feasibility of the motor system. The meteorological observation model of 400 kgf class motor is under development and the development of microgravity experiment class of 1.5-2 tonf motor will follow subsequently. The authors plan to complete the development of the 400 kgf class motor for meteorological observation model by the end of FY2005.

  11. Ceramic matrix composites for rocket engine turbine applications

    NASA Technical Reports Server (NTRS)

    Herbell, Thomas P.; Eckel, Andrew J.

    1992-01-01

    A program to establish the potential for introducing fiber reinforced ceramic matrix composites (FRCMC) in future rocket engine turbopumps was instituted in 1987. A brief summary of the overall program (both contract and in-house research) is presented. Tests at NASA Lewis include thermal upshocks in a hydrogen/oxygen test rig capable of generating heating rates up to 2500 C/sec. Post thermal upshock exposure evaluation includes the measurement of residual strength and failure analysis. Test results for monolithic ceramics and several FRCMC are presented. Hydrogen compatibility was assessed by isothermal exposure of monolithic ceramics in high temperature gaseous hydrogen plus water vapor.

  12. GASEOUS SCINTILLATION COUNTER

    DOEpatents

    Eggler, C.; Huddleston, C.M.

    1959-04-28

    A gaseous excitation counter for detecting the presence amd measuring the energy of subatomic particles and electromagnetic radiation is described. The counter includes a gas-tight chamber filled with an elemental gas capable of producing ultra-violet excitation quanta when irradiated with subatomic particles and electromagnetic radiation. The gas has less than one in a thousand parts ultra-violet absorbing contamination. When nuclear radiation ps present the ultra-violet light produced by the gas strikes a fluorescent material within the counter, responsive to produce visible excitation quanta, and photo-sensitive counting means detect the visible emission.

  13. Rocket + Science = Dialogue

    NASA Technical Reports Server (NTRS)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

    It's a cliche that rocket engineers and space scientists don t see eye-to-eye. That goes double for rocket engineers working on human spaceflight and scientists working on space telescopes and planetary probes. They work fundamentally different problems but often feel that they are competing for the same pot of money. Put the two groups together for a weekend, and the results could be unscientific or perhaps combustible. Fortunately, that wasn't the case when NASA put heavy lift launch vehicle designers together with astronomers and planetary scientists for two weekend workshops in 2008. The goal was to bring the top people from both groups together to see how the mass and volume capabilities of NASA's Ares V heavy lift launch vehicle could benefit the science community. Ares V is part of NASA's Constellation Program for resuming human exploration beyond low Earth orbit, starting with missions to the Moon. In the current mission scenario, Ares V launches a lunar lander into Earth orbit. A smaller Ares I rocket launches the Orion crew vehicle with up to four astronauts. Orion docks with the lander, attached to the Ares V Earth departure stage. The stage fires its engine to send the mated spacecraft to the Moon. Standing 360 feet high and weighing 7.4 million pounds, NASA's new heavy lifter will be bigger than the 1960s-era Saturn V. It can launch almost 60 percent more payload to translunar insertion together with the Ares I and 35 percent more mass to low Earth orbit than the Saturn V. This super-sized capability is, in short, designed to send more people to more places to do more things than the six Apollo missions.

  14. Rocket Engine Thrust Chamber Assembly

    NASA Technical Reports Server (NTRS)

    Cornelius, Charles S. (Inventor); Counts, Richard H. (Inventor); Myers, W. Neill (Inventor); Lackey, Jeffrey D. (Inventor); Peters, Warren (Inventor); Shadoan, Michael (Inventor); Sparks, David L. (Inventor); Lawrence, Timothy W. (Inventor)

    2001-01-01

    A thrust chamber assembly for liquid fueled rocket engines and the method of making it wherein a two-piece mandrel having the configuration of an assembly having a combustion chamber portion connected to a nozzle portion through a throat portion is wrapped with a silica tape saturated with a phenolic resin, the tape extending along the mandrel and covering the combustion chamber portion of the mandrel to the throat portion. The width of the tape is positioned at an angle of 30 to 50 deg. to the axis of the mandrel such that one edge of the tape contacts the mandrel while the other edge is spaced from the mandrel. The phenolic in the tape is cured and the end of the wrap is machined to provide a frusto-conical surface extending at an angle of 15 to 30 deg. with respect to the axis of the mandrel for starting a second wrap on the mandrel to cover the throat portion. The remainder of the mandrel is wrapped with a third silica tape having its width positioned at a angle of 5 to 20 deg. from the axis of the mandrel. The resin in the third tape is cured and the assembly is machined to provide a smooth outer surface. The entire assembly is then wrapped with a tow of graphite fibers wetted with an epoxy resin and, after the epoxy resin is cured, the graphite is machined to final dimensions.

  15. Coal-Fired Rocket Engine

    NASA Technical Reports Server (NTRS)

    Anderson, Floyd A.

    1987-01-01

    Brief report describes concept for coal-burning hybrid rocket engine. Proposed engine carries larger payload, burns more cleanly, and safer to manufacture and handle than conventional solid-propellant rockets. Thrust changeable in flight, and stops and starts on demand.

  16. Otrag rocket experiments in Africa

    NASA Technical Reports Server (NTRS)

    1978-01-01

    West German rocket manufacturers are testing their products in Zaire. Hundreds of pipes (12 m x 80 cm) are bundled together inside the test missiles, which are fired into Zaire's prairie. The reactions of neighboring nations, as well as leading countries of the world, are presented concerning the rocket tests.

  17. Fluid thrust control system. [for liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Howell, W. L.; Jansen, H. B.; Lehmann, E. N. (Inventor)

    1968-01-01

    A pure fluid thrust control system is described for a pump-fed, regeneratively cooled liquid propellant rocket engine. A proportional fluid amplifier and a bistable fluid amplifier control overshoot in the starting of the engine and take it to a predetermined thrust. An ejector type pump is provided in the line between the liquid hydrogen rocket nozzle heat exchanger and the turbine driving the fuel pump to aid in bringing the fluid at this point back into the regular system when it is not bypassed. The thrust control system is intended to function in environments too severe for mechanical controls.

  18. Computational neutronic analysis of the nuclear vapor thermal rocket engine

    SciTech Connect

    Dugan, E.T.; Watanabe, Y.; Kuras, S.; Maya, I.; Diaz, N.J. )

    1992-01-01

    Calculational procedures and results are presented for the neutronic analysis of the Nuclear Vapor Thermal Reactor (NVTR) rocket engine. The NVTR, in a rocket engine, uses modified NERVA geometry and systems with the solid fuel replaced by highly enriched (>85%) uranium tetrafluoride (UF[sub 4]) vapor. In the NVTR, the hydrogen propellant is the primary coolant, is physically separated from the UF[sub 4] vapor (which is not circulated), is maintained at high pressure (50 to 100 atm), and exits the core at 3100 to 3500 K.

  19. Preliminary Studies on a Small-Scale Single-Tube Pulse Detonation Rocket Prototype

    NASA Astrophysics Data System (ADS)

    Wang, Ke; Fan, Wei; Yan, Yu; Jin, Le

    2013-06-01

    As a new concept propulsion system, the pulse detonation engine has received extensive concerns from all over the world in the past few years. With oxidizer on board, it operates as a rocket engine which is known as pulse detonation rocket engine. In this study, a rocket model powered by a single-tube pulse detonation rocket engine was fabricated to demonstrate and validate whether or not it could operate stably and reliably independently. The single-tube pulse detonation rocket prototype consisted of a wireless control unit, three tanks for oxidizer, fuel and purge gas, various valves and a detonation tube. With compact design, the pulse detonation rocket prototype had an outer diameter of 260 mm and a length of 2200 mm. Oxygen, liquid aviation kerosene and nitrogen were utilized as oxidizer, fuel and purge gas, respectively. Operation tests were carried out to obtain proper operating conditions for the pulse detonation rocket prototype first, and then sliding test was conducted. It was concluded that the pulse detonation rocket prototype could operate stably and reliably. The generated thrust was estimated and compared with theoretical value.

  20. RECENT ACTIVITIES AT THE CENTER FOR SPACE NUCLEAR RESEARCH FOR DEVELOPING NUCLEAR THERMAL ROCKETS

    SciTech Connect

    Robert C. O'Brien

    2001-09-01

    Nuclear power has been considered for space applications since the 1960s. Between 1955 and 1972 the US built and tested over twenty nuclear reactors/ rocket-engines in the Rover/NERVA programs. However, changes in environmental laws may make the redevelopment of the nuclear rocket more difficult. Recent advances in fuel fabrication and testing options indicate that a nuclear rocket with a fuel form significantly different from NERVA may be needed to ensure public support. The Center for Space Nuclear Research (CSNR) is pursuing development of tungsten based fuels for use in a NTR, for a surface power reactor, and to encapsulate radioisotope power sources. The CSNR Summer Fellows program has investigated the feasibility of several missions enabled by the NTR. The potential mission benefits of a nuclear rocket, historical achievements of the previous programs, and recent investigations into alternatives in design and materials for future systems will be discussed.

  1. Rhenium Rocket Manufacturing Technology

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The NASA Lewis Research Center's On-Board Propulsion Branch has a research and technology program to develop high-temperature (2200 C), iridium-coated rhenium rocket chamber materials for radiation-cooled rockets in satellite propulsion systems. Although successful material demonstrations have gained much industry interest, acceptance of the technology has been hindered by a lack of demonstrated joining technologies and a sparse materials property data base. To alleviate these concerns, we fabricated rhenium to C-103 alloy joints by three methods: explosive bonding, diffusion bonding, and brazing. The joints were tested by simulating their incorporation into a structure by welding and by simulating high-temperature operation. Test results show that the shear strength of the joints degrades with welding and elevated temperature operation but that it is adequate for the application. Rhenium is known to form brittle intermetallics with a number of elements, and this phenomena is suspected to cause the strength degradation. Further bonding tests with a tantalum diffusion barrier between the rhenium and C-103 is planned to prevent the formation of brittle intermetallics.

  2. Advanced materials for radiation-cooled rockets

    NASA Astrophysics Data System (ADS)

    Reed, Brian; Biaglow, James; Schneider, Steven

    1993-11-01

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

  3. Advanced materials for radiation-cooled rockets

    NASA Technical Reports Server (NTRS)

    Reed, Brian; Biaglow, James; Schneider, Steven

    1993-01-01

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

  4. Supercomputer modeling of hydrogen combustion in rocket engines

    NASA Astrophysics Data System (ADS)

    Betelin, V. B.; Nikitin, V. F.; Altukhov, D. I.; Dushin, V. R.; Koo, Jaye

    2013-08-01

    Hydrogen being an ecological fuel is very attractive now for rocket engines designers. However, peculiarities of hydrogen combustion kinetics, the presence of zones of inverse dependence of reaction rate on pressure, etc. prevents from using hydrogen engines in all stages not being supported by other types of engines, which often brings the ecological gains back to zero from using hydrogen. Computer aided design of new effective and clean hydrogen engines needs mathematical tools for supercomputer modeling of hydrogen-oxygen components mixing and combustion in rocket engines. The paper presents the results of developing verification and validation of mathematical model making it possible to simulate unsteady processes of ignition and combustion in rocket engines.

  5. The Solid Rocket Booster Auxiliary Power Unit: Meeting the Challenge

    NASA Technical Reports Server (NTRS)

    Hughes, R. W.

    1985-01-01

    The thrust vector control systems of the solid rocket boosters are turbine-powered, electrically controlled hydraulic systems which function through hydraulic actuators to gimbal the nozzles of the solid rocket boosters and provide vehicle steering for the Space Shuttle. Turbine power for the thrust vector control systems is provided through hydrazine fueled auxiliary power units which drive the hydraulic pumps. The solid rocket booster auxiliary power unit resulted from trade studies which indicated significant advantages would result if an existing engine could be found to meet the program goal of 20 missions reusability and adapted to meet the seawater environments associated with ocean landings. During its maturation, the auxiliary power unit underwent many design iterations and provided its flight worthiness through full qualification programs both as a component and as part of the thrust vector control system. More significant, the auxiliary power unit has successfully completed six Shuttle missions.

  6. Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

    NASA Technical Reports Server (NTRS)

    Shafirovich, E. IA.; Shiriaev, A. A.; Goldshleger, U. I.

    1993-01-01

    A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

  7. Nuclear Thermal Rocket Element Environmental Simulator (NTREES) Upgrade Activities

    NASA Technical Reports Server (NTRS)

    Emrich, William J. Jr.; Moran, Robert P.; Pearson, J. Boise

    2012-01-01

    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). This device can simulate 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 so as to accurately reproduce the temperatures and heat fluxes which would normally occur as a result of nuclear fission and would be exposed to flowing hydrogen. Initial testing of a somewhat prototypical fuel element has been successfully performed in NTREES and the facility has now been shutdown to allow for an extensive reconfiguration of the facility which will result in a significant upgrade in its capabilities

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

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. This... 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. 14 CFR 34.71 - Compliance with gaseous emission standards.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. This... 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) §...

  10. 14 CFR 34.71 - Compliance with gaseous emission standards.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. This... 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) §...

  11. Combustion of metal agglomerates in a solid rocket core flow

    NASA Astrophysics Data System (ADS)

    Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

    2013-12-01

    The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

  12. Rocket/launcher structural dynamics

    NASA Technical Reports Server (NTRS)

    Ferragut, N. J.

    1976-01-01

    The equations of motion describing the interactions between a rocket and a launcher were derived using Lagrange's Equation. A rocket launching was simulated. The motions of both the rocket and the launcher can be considered in detail. The model contains flexible elements and rigid elements. The rigid elements (masses) were judiciously utilized to simplify the derivation of the equations. The advantages of simultaneous shoe release were illustrated. Also, the loading history of the interstage structure of a boosted configuration was determined. The equations shown in this analysis could be used as a design tool during the modification of old launchers and the design of new launchers.

  13. Exergy Analysis of Rocket Systems

    NASA Technical Reports Server (NTRS)

    Gilbert, Andrew; Mesmer, Bryan; Watson, Michael D.

    2015-01-01

    Exergy is defined as the useful work available from a system in a specified environment. Exergy analysis allows for comparison between different system designs, and allows for comparison of subsystem efficiencies within system designs. The proposed paper explores the relationship between the fundamental rocket equation and an exergy balance equation. A previously derived exergy equation related to rocket systems is investigated, and a higher fidelity analysis will be derived. The exergy assessments will enable informed, value-based decision making when comparing alternative rocket system designs, and will allow the most efficient configuration among candidate configurations to be determined.

  14. Dynamic characterization of solid rockets

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The structural dynamics of solid rockets in-general was studied. A review is given of the modes of vibration and bending that can exist for a solid propellant rocket, and a NASTRAN computer model is included. Also studied were the dynamic properties of a solid propellant, polybutadiene-acrylic acid-acrylonitrile terpolymer, which may be used in the space shuttle rocket booster. The theory of viscoelastic materials (i.e, Poisson's ratio) was employed in describing the dynamic properties of the propellant. These studies were performed for an eventual booster stage development program for the space shuttle.

  15. 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

  16. A case for Mars: A case for nuclear thermal rockets

    SciTech Connect

    Neuman, J.E.; Van Haaften, D.H.; Madsen, W.W.

    1990-01-01

    It is now possible to make general comparisons of candidate propulsion systems for human exploration of Mars. Preliminary review indicates that the propulsion system most likely to meet all mission requirements is the Nuclear Thermal Rocket (NTR). Advanced cryogenic chemical propulsion systems achieve a maximum specific impulse (Isp) of about 470 seconds. The Nuclear Engine for Rocket Vehicle Application (NERVA) program of the 1960's built engines with Isp's of about 825 seconds. Performance of an NTR depends on achievable materials temperatures, but materials has progressed significantly since the 1960's. Also, some of the current research undertaken to improve chemical rocket performance, such as aerobraking or schemes to minify payload, applies to an NTR as well, although it is not essential. The NTR is reusable, and can be developed into a complete space transportation system. Only 3--4% of the nuclear fuel would be used in a Mars mission, and an engine can be used until about 40% of the fuel is expended. Nuclear thermal rockets can take mankind to the moon, to Mars, and beyond, but development must begin now. There is potential for orderly growth into nuclear concepts far beyond NERVA. Using chemical propulsion for lunar missions and delaying NTR development will only result in higher costs and delayed or cancelled Mars missions.

  17. World Data Center A (rockets and satellites) catalogue of data. Volume 1, part A: Sounding rockets

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A cumulative listing of all scientifically successful rockets that have been identified from various sources is presented. The listing starts with the V-2 rocket launched on 7 March 1947 and contains all rockets identified up to 31 December 1971.

  18. Navigating the Rockets Educator Guide

    NASA Video Gallery

    In this brief video overview, learn how to navigate the Rockets Educator Guide. Get a glimpse of the resources available in the guide, including a pictorial history, an overview of the physics cont...

  19. Small Solid Rocket Motor Test

    NASA Video Gallery

    It was three-two-one to brilliant fire as NASA's Marshall Space Flight Center tested a small solid rocket motor designed to mimic NASA's Space Launch System booster. The Mar. 14 test provides a qui...

  20. Studies of the exhaust products from solid propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Dawbarn, R.; Kinslow, M.

    1976-01-01

    This study was undertaken to determine the feasibility of conducting environmental chamber tests on the physical processes which occur when a solid rocket motor exhaust mixes with the ambient atmosphere. Of particular interest was the interaction between hydrogen chloride, aluminum oxide, and water vapor. The program consisted of three phases: (1) building a small rocket motor and using it to provide the exhaust species in a controlled environment; (2) evaluating instruments used to detect and measure HCl concentrations and if possible determining whether the HCl existed in the gaseous state or as an acid aerosol; (3) monitoring a series of 6.4-percent scale space shuttle motor tests and comparing the results to the environmental chamber studies. Eighteen firings were conducted in an environmental chamber with the initial ambient relative humidity set at values from 29 to 100 percent. Two additional firings were made in a large shed, and four were made on an open concrete apron. Six test firings at MSFC were monitored, and the ground level concentrations are reported. Evidence is presented which shows that the larger Al2O3 (5 to 50 micrometers) particles from the rocket motor can act as condensation nuclei. Under appropriate ambient conditions where there is sufficient water vapor this results in the formation of an acid aerosol. Droplets of this acid were detected both in the environmental chamber and in the scaled shuttle engine tests.