Science.gov

Sample records for rocket propellants

  1. 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.

  2. Casting propellant in rocket engine

    NASA Technical Reports Server (NTRS)

    Roach, J. E.; Froehling, S. C. (Inventor)

    1976-01-01

    A method is described for casting a solid propellant in the casing of a rocket engine having a continuous wall with a single opening which is formed by leaves of a material which melt at a temperature of the propellant and with curved edges concentric to the curvature of the spherical casing. The leaves are inserted into the spherical casing through the opening forming a core having a greater width than the width of the single opening and with curved peripheral edges. The cast propellant forms a solid mass and then heated to melt the leaves and provide a central opening with radial projecting flutes.

  3. Liquid propellant rocket combustion instability

    NASA Technical Reports Server (NTRS)

    Harrje, D. T.

    1972-01-01

    The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

  4. Environmentally compatible solid rocket propellants

    NASA Technical Reports Server (NTRS)

    Jacox, James L.; Bradford, Daniel J.

    1995-01-01

    Hercules' clean propellant development research is exploring three major types of clean propellant: (1) chloride-free formulations (no chlorine containing ingredients), being developed on the Clean Propellant Development and Demonstration (CPDD) contract sponsored by Phillips Laboratory, Edwards Air Force Base, CA; (2) low HCl scavenged formulations (HCl-scavenger added to propellant oxidized with ammonium perchlorate (AP)); and (3) low HCl formulations oxidized with a combination of AN and AP (with or without an HCl scavenger) to provide a significant reduction (relative to current solid rocket boosters) in exhaust HCl. These propellants provide performance approaching that of current systems, with less than 2 percent HCl in the exhaust, a significant reduction (greater than or equal to 70 percent) in exhaust HCl levels. Excellent processing, safety, and mechanical properties were achieved using only readily available, low cost ingredients. Two formulations, a sodium nitrate (NaNO3) scavenged HTPB and a chloride-free hydroxy terminated polyether (HTPE) propellant, were characterized for ballistic, mechanical, and rheological properties. In addition, the hazards properties were demonstrated to provide two families of class 1.3, 'zero-card' propellants. Further characterization is planned which includes demonstration of ballistic tailorability in subscale (one to 70 pound) motors over the range of burn rates required for retrofit into current Hercules space booster designs (Titan 4 SRMU and Delta 2 GEM).

  5. Feasibility of rocket propellant production on Mars

    NASA Technical Reports Server (NTRS)

    Ash, R. L.; Dowler, W. L.; Varsi, G.

    1978-01-01

    In situ production of rocket propellant to reduce landed mass requirements for Mars return missions has been investigated. The analysis has shown that a system which utilizes atmospheric carbon dioxide and soil moisture to produce liquid methane-oxygen propellant requires a landed mass which is less than half the mass of the ascent vehicle it produces.

  6. Rocket thrust variation with foamed liquid propellants

    NASA Technical Reports Server (NTRS)

    Morrell, G

    1957-01-01

    An analysis is presented on a method for varying rocket thrust by varying the bulk density of the propellants. This density variation was accomplished by uniformly dispersing an inert, insoluble gas in the liquid propellants. Only qualitative agreement with theory was obtained from preliminary experiments with a 1000-pound-thrust ammonia - nitric acid rocket engine; the required experimental gas-flow rates were two to six times greater than those predicted by theory. It was demonstrated, however, that this method of rocket-thrust variation is feasible.

  7. A miniature solid propellant rocket motor

    SciTech Connect

    Grubelich, M.C.; Hagan, M.; Mulligan, E.

    1997-08-01

    A miniature solid-propellant rocket motor has been developed to impart a specific motion to an object deployed in space. This rocket motor effectively eliminated the need for a cold-gas thruster system or mechanical spin-up system. A low-energy igniter, an XMC4397, employing a semiconductor bridge was used to ignite the rocket motor. The rocket motor was ground-tested in a vacuum tank to verify predicted space performance and successfully flown in a Sandia National Laboratories flight vehicle program.

  8. Acceleration effects in solid propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Langhenry, M. T.

    1986-01-01

    The performance variations due to acceleration loads imposed on spinning solid propellant rocket motors are investigated. The four potentially most significant modes of acceleration-induced phenomena are identified from a study of the literature and modeled. The four modes are a mechanical mode which deals with deformations of the propellant and case: a thermodynamic mode which covers acceleration-induced combustion phenomena; a stress mode which covers the stressed propellant's effect on burn rate; and a gas dynamic mode which deals with changes in gas flow in the chamber and through the nozzle. Simplified models of each mode are developed or taken from the literature and are added to an internal ballistics evaluation computer program. The resulting analysis is the first to include all of the modes. In order to do this an original analysis of the mechanical and stress modes was necessary. However, the analysis shows that the stress mode is not important for the circular perforated grains studied. The other effects are shown to have a significant influence on solid rocket motor performance. The magnitude of the different mode effects are such that one may not be ignored over the others as has been done in the past. The results of the analysis are compared to published rocket motor data. The comparisons indicate an erosive burning effect that is a function of spin rate. A qualitative explanation of the erosive effect is presented.

  9. Nuclear thermal rockets using indigenous extraterrestrial propellants

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert M.

    1990-01-01

    A preliminary examination of a concept for a Mars and outer solar system exploratory vehicle is presented. Propulsion is provided by utilizing a nuclear thermal reactor to heat a propellant volatile indigenous to the destination world to form a high thrust rocket exhaust. Candidate propellants, whose performance, materials compatibility, and ease of acquisition are examined and include carbon dioxide, water, methane, nitrogen, carbon monoxide, and argon. Ballistics and winged supersonic configurations are discussed. It is shown that the use of this method of propulsion potentially offers high payoff to a manned Mars mission. This is accomplished by sharply reducing the initial mission mass required in low earth orbit, and by providing Mars explorers with greatly enhanced mobility in traveling about the planet through the use of a vehicle that can refuel itself each time it lands. Thus, the nuclear landing craft is utilized in combination with a hydrogen-fueled nuclear-thermal interplanetary launch. By utilizing such a system in the outer solar system, a low level aerial reconnaissance of Titan combined with a multiple sample return from nearly every satellite of Saturn can be accomplished in a single launch of a Titan 4 or the Space Transportation System (STS). Similarly a multiple sample return from Callisto, Ganymede, and Europa can also be accomplished in one launch of a Titan 4 or the STS.

  10. 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

  11. Solid propellant processing factor in rocket motor design

    NASA Technical Reports Server (NTRS)

    1971-01-01

    The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

  12. Propellant-Flow-Actuated Rocket Engine Igniter

    NASA Technical Reports Server (NTRS)

    Wollen, Mark

    2013-01-01

    A rocket engine igniter has been created that uses a pneumatically driven hammer that, by specialized geometry, is induced into an oscillatory state that can be used to either repeatedly impact a piezoelectric crystal with sufficient force to generate a spark capable of initiating combustion, or can be used with any other system capable of generating a spark from direct oscillatory motion. This innovation uses the energy of flowing gaseous propellant, which by means of pressure differentials and kinetic motion, causes a hammer object to oscillate. The concept works by mass flows being induced through orifices on both sides of a cylindrical tube with one or more vent paths. As the mass flow enters the chamber, the pressure differential is caused because the hammer object is supplied with flow on one side and the other side is opened with access to the vent path. The object then crosses the vent opening and begins to slow because the pressure differential across the ball reverses due to the geometry in the tube. Eventually, the object stops because of the increasing pressure differential on the object until all of the kinetic energy has been transferred to the gas via compression. This is the point where the object reverses direction because of the pressure differential. This behavior excites a piezoelectric crystal via direct impact from the hammer object. The hammer strikes a piezoelectric crystal, then reverses direction, and the resultant high voltage created from the crystal is transferred via an electrode to a spark gap in the ignition zone, thereby providing a spark to ignite the engine. Magnets, or other retention methods, might be employed to favorably position the hammer object prior to start, but are not necessary to maintain the oscillatory behavior. Various manifestations of the igniter have been developed and tested to improve device efficiency, and some improved designs are capable of operation at gas flow rates of a fraction of a gram per second (0

  13. Coated oxidizers for combustion stability in solid-propellant rockets

    NASA Technical Reports Server (NTRS)

    Helmy, A. M.; Ramohalli, K. N. R.

    1985-01-01

    Experiments are conducted in a laboratory-scale (6.25-cm diameter) end-burning rocket motor with state-of-the-art, ammonium perchlorate hydroxy-terminated polybutadiene (HTPB), nonmetallized propellants. The concept of tailoring the stability characteristics with a small amount (less than 1 percent by weight) of COATING on the oxidizer is explored. The thermal degradation characteristics of the coat chemical are deduced through theoretical arguments on thermal diffusivity of the composite material (propellant). Several candidate coats are selected and propellants are cast. These propellants (with coated oxidizers) are fired in a laboratory-scale end-burning rocket motor, and real-time pressure histories are recorded. The control propellant (with no coating) is also tested for comparison. The uniformity of the coating, confirmed by SEM pictures and BET adsorption measurements, is thought to be an advance in technology. The frequency of bulk mode instability (BMI), the pressure fluctuation amplitudes, and stability boundaries are correlated with parameters related to the characteristic length (L-asterisk) of the rocket motor. The coated oxidizer propellants, in general, display greater combustion stability than the control (state-of-the-art). The correlations of the various parameters are thought to be new to a field filled with much uncertainty.

  14. Viability of Bacillus subtilis Spores in Rocket Propellants

    PubMed Central

    Godding, Rogene M.; Lynch, Victoria H.

    1965-01-01

    The sporicidal activity of components used in liquid and solid rocket propellants was tested by use of spores of Bacillus subtilis dried on powdered glass. Liquid propellant ingredients tested were N2O4, monomethylhydrazine and 1,1-dimethylhydrazine. N2O4 was immediately sporicidal; the hydrazines were effective within several days. Solid propellants consisted of ammonium perchlorate in combination with epoxy resin (EPON 828), tris-1-(2-methyl) aziridinyl phosphine oxide, bis-1-(2-methyl) aziridinyl phenylphosphine oxide, and three modified polybutadiene polymers. There was no indication of appreciable sporicidal activity of these components. PMID:14264838

  15. Motion Analysis of a Rocket-Propelled Truck.

    ERIC Educational Resources Information Center

    Hitt, Darren L.; Lowe, Mary L.

    1996-01-01

    Describes an experiment to study the motion of a rocket-propelled vehicle over the entire duration of the engine burn using a video system with a frame-by-frame playback and a Sonic Ranger for ultrasonic position movements. Enables students to study the impulse-momentum principle and the effects of a time-varying force. (JRH)

  16. Dragonfly directional sensor versus rocket-propelled grenades

    NASA Astrophysics Data System (ADS)

    Geary, Joseph; Blackwell, Lisa

    2015-02-01

    The Dragonfly directional sensor was deployed at the Army's Yuma Proving Grounds for preliminary field tests against rocket-propelled grenades. This wide-field (nonimaging) sensor's purpose was to angularly locate the latter's launch plume. These tests successfully demonstrated proof-of-concept.

  17. 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.

  18. Rocket Propellants Engine Design/Operations/Validation

    NASA Technical Reports Server (NTRS)

    Monk, Jan C.

    2002-01-01

    Lockheed Martin Astronautics Operations (LMA) was competitively awarded a contract May 21, 2001 for next generation launch system architecture definition and technology maturation. The Second Generation Launch Vehicle Program objectives include reducing the technical and programmatic risk of proceeding to full scale development of the system by establishing requirements for the next generation launch system and maturing critical technologies needed by the system. LMA will conduct analyses and trades to optimize the architecture ETO elements including configuration, conceptual designs, and preliminary operations definition. To fully understand the engine and propellant trades were conducted by LMA to yield the optimized architecture system from the operability, reliability, safety, and cost perspectives. A government/industry team addressed the required trade studies, the parameters and weighting factors, and the most critical trades were addressed. This report summarizes the participation of JCM Consulting, Inc. in the propellant trade study.

  19. 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.

  20. High-pressure burning rate studies of solid rocket propellants

    NASA Astrophysics Data System (ADS)

    Atwood, A. I.; Ford, K. P.; Wheeler, C. J.

    2013-03-01

    Increased rocket motor performance is a major driver in the development of solid rocket propellant formulations for chemical propulsion systems. The use of increased operating pressure is an option to improve performance potentially without the cost of reformulation. A technique has been developed to obtain burning rate data across a range of pressures from ambient to 345 MPa. The technique combines the use of a low loading density combustion bomb with a high loading density closed bomb technique. A series of nine ammonium perchlorate (AP) based propellants were used to demonstrate the use of the technique, and the results were compared to the neat AP burning rate "barrier". The effect of plasticizer, oxidizer particle size, catalyst, and binder type were investigated.

  1. Characterization of rocket propellant combustion products

    SciTech Connect

    Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.

    1991-12-09

    The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

  2. 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.

  3. Erosive burning research. [for solid-propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Strand, L.; Yang, L. C.; Nguyen, M. H.; Cohen, N. S.

    1986-01-01

    A status report is given on the results for the completed tests in a series of motor firings being carried out to measure the effects of the parameters that are considered to most strongly influence the scaling to larger rocket motor sizes of the transition to/or threshold conditions for erosive burning rate augmentation. Propellant burning rates at locations along the axis of the test motors are measured with a newly developed plasma capacitance gauge technique. The measured results are compared with erosive-burning predictions from a supporting ballistics analysis. The completed motor firings have successfully demonstrated response to the designed test variables. The trends with varying propellant burning rate, chamber pressure, and mass flow rate are consistent with existing results, but no pronounced effect of surface roughness has been observed. Rather, the influence of propellant oxidizer particle size on erosive burning is through its effect on the base, no-corssflow burning rate.

  4. 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.

  5. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 8 2010-07-01 2010-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor...

  6. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 8 2011-07-01 2011-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor...

  7. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 9 2013-07-01 2013-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor...

  8. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 9 2012-07-01 2012-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor...

  9. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 9 2014-07-01 2014-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor...

  10. On the history of the development of solid-propellant rockets in the Soviet Union

    NASA Technical Reports Server (NTRS)

    Pobedonostsev, Y. A.

    1977-01-01

    Pre-World War II Soviet solid-propellant rocket technology is reviewed. Research and development regarding solid composite preparations of pyroxyline TNT powder is described, as well as early work on rocket loading calculations, problems of flight stability, and aircraft rocket launching and ground rocket launching capabilities.

  11. Development and Test of a Rocket Engine Using Environmentally Friendly Propellants

    NASA Technical Reports Server (NTRS)

    Webster, Kristi

    2009-01-01

    Develop and test a rocket engine that operates on environmentally friendly propellants; Liquid Oxygen (LOX) and Liquid Methane (LCH4). Due to modifications the rocket engine designed last summer (KJ_REX) is not the same rocket thruster tested this summer, but very similar. The new modified rocket thruster was built for NASA by Orion Propulsion Inc. (OPI), Huntsville, AL.

  12. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    Thomas, Matthew E.; Bossard, John A.; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

    2001-01-01

    The fiber optically coupled laser ignition approach summarized is under consideration for use in igniting bi-propellant rocket thrust chambers. This laser ignition approach is based on a novel dual pulse format capable of effectively increasing laser generated plasma life times up to 1000 % over conventional laser ignition methods. In the dual-pulse format tinder consideration here an initial laser pulse is used to generate a small plasma kernel. A second laser pulse that effectively irradiates the plasma kernel follows this pulse. Energy transfer into the kernel is much more efficient because of its absorption characteristics thereby allowing the kernel to develop into a much more effective ignition source for subsequent combustion processes. In this research effort both single and dual-pulse formats were evaluated in a small testbed rocket thrust chamber. The rocket chamber was designed to evaluate several bipropellant combinations. Optical access to the chamber was provided through small sapphire windows. Test results from gaseous oxygen (GOx) and RP-1 propellants are presented here. Several variables were evaluated during the test program, including spark location, pulse timing, and relative pulse energy. These variables were evaluated in an effort to identify the conditions in which laser ignition of bi-propellants is feasible. Preliminary results and analysis indicate that this laser ignition approach may provide superior ignition performance relative to squib and torch igniters, while simultaneously eliminating some of the logistical issues associated with these systems. Further research focused on enhancing the system robustness, multiplexing, and window durability/cleaning and fiber optic enhancements is in progress.

  13. 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.

  14. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2007-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  15. Scaling of Performance in Liquid Propellant Rocket Engine Combustion Devices

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2008-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  16. Terminology and assessment methods of solid propellant rocket exhaust signatures

    NASA Astrophysics Data System (ADS)

    1993-02-01

    The Propulsion and Energetics Panel's Specialists' Meeting in autumn 1985 on Smokeless Propellants demonstrated that no common standard was available in this field and that the lack of common understanding led to misunderstanding amongst the NATO community. After some preparatory discussion, the Panel, therefore, formed Working Group Number 21 with the objectives of defining methods for the assessment of rocket motor exhaust optical properties in the visible and in the infrared range, and of recommending a terminology based on quantitative criteria. The Working Group discussed the subject in a total of eight sessions and prepared this Advisory Report. Following an Introduction and Summary there are six chapters, commencing with an Overview and continuing with Propellant Smoke Classification, Plume Primary Smoke, Plume Secondary Smoke, Plume Radiation and Plume Microwave Properties. In most cases, the conclusions and recommendations follow the chapters and are not repeated at the end of the report.

  17. Materials Problems in Chemical Liquid-Propellant Rocket Systems

    NASA Technical Reports Server (NTRS)

    Gilbert, L. L.

    1959-01-01

    With the advent of the space age, new adjustments in technical thinking and engineering experience are necessary. There is an increasing and extensive interest in the utilization of materials for components to be used at temperatures ranging from -423 to over 3500 deg F. This paper presents a description of the materials problems associated with the various components of chemical liquid rocket systems. These components include cooled and uncooled thrust chambers, injectors, turbine drive systems, propellant tanks, and cryogenic propellant containers. In addition to materials limitations associated with these components, suggested research approaches for improving materials properties are made. Materials such as high-temperature alloys, cermets, carbides, nonferrous alloys, plastics, refractory metals, and porous materials are considered.

  18. SRM (Solid Rocket Motor) propellant and polymer materials structural modeling

    NASA Technical Reports Server (NTRS)

    Moore, Carleton J.

    1988-01-01

    The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

  19. Boundary cooled rocket engines for space storable propellants

    NASA Technical Reports Server (NTRS)

    Kesselring, R. C.; Mcfarland, B. L.; Knight, R. M.; Gurnitz, R. N.

    1972-01-01

    An evaluation of an existing analytical heat transfer model was made to develop the technology of boundary film/conduction cooled rocket thrust chambers to the space storable propellant combination oxygen difluoride/diborane. Critical design parameters were identified and their importance determined. Test reduction methods were developed to enable data obtained from short duration hot firings with a thin walled (calorimeter) chamber to be used quantitatively evaluate the heat absorbing capability of the vapor film. The modification of the existing like-doublet injector was based on the results obtained from the calorimeter firings.

  20. 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.

  1. On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

    NASA Technical Reports Server (NTRS)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

  2. Computational Thermochemistry of Jet Fuels and Rocket Propellants

    NASA Technical Reports Server (NTRS)

    Crawford, T. Daniel

    2002-01-01

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

  3. Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

    NASA Technical Reports Server (NTRS)

    Casiano, Matthew; Hulka, James; Yang, Virog

    2009-01-01

    Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

  4. Solid-propellant rocket motor ballistic performance variation analyses

    NASA Technical Reports Server (NTRS)

    Sforzini, R. H.; Foster, W. A., Jr.

    1975-01-01

    Results are presented of research aimed at improving the assessment of off-nominal internal ballistic performance including tailoff and thrust imbalance of two large solid-rocket motors (SRMs) firing in parallel. Previous analyses using the Monte Carlo technique were refined to permit evaluation of the effects of radial and circumferential propellant temperature gradients. Sample evaluations of the effect of the temperature gradients are presented. A separate theoretical investigation of the effect of strain rate on the burning rate of propellant indicates that the thermoelastic coupling may cause substantial variations in burning rate during highly transient operating conditions. The Monte Carlo approach was also modified to permit the effects on performance of variation in the characteristics between lots of propellants and other materials to be evaluated. This permits the variabilities for the total SRM population to be determined. A sample case shows, however, that the effect of these between-lot variations on thrust imbalances within pairs of SRMs is minor in compariosn to the effect of the within-lot variations. The revised Monte Carlo and design analysis computer programs along with instructions including format requirements for preparation of input data and illustrative examples are presented.

  5. Theoretical performance of lithium and fluorine as a rocket propellant

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Huff, Vearl N

    1951-01-01

    Theoretical performance for liquid lithium and liquid fluorine as a rocket propellant was calculated with assumptions both of equilibrium and frozen composition during expansion. Parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, composition, mean molecular weight, characteristic velocity, coefficient of thrust, and ratio of nozzle-exit area to throat area. For chamber pressure of 300 pounds per square inch absolute and expansion to 1 atmosphere, the maximum equilibrium specific impulse calculated was 335.5 pound-seconds per pound. The effect of ionization on calculated performance was shown to be negligible by comparison of values of various parameters calculated both with and without ionized products of combustion.

  6. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James

    2008-01-01

    The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

  7. Lidar measurements of solid rocket propellant fire particle plumes.

    PubMed

    Brown, David M; Brown, Andrea M; Willitsford, Adam H; Dinello-Fass, Ryan; Airola, Marc B; Siegrist, Karen M; Thomas, Michael E; Chang, Yale

    2016-06-10

    This paper presents the first, to our knowledge, direct measurement of aerosol produced by an aluminized solid rocket propellant (SRP) fire on the ground. Such fires produce aluminum oxide particles small enough to loft high into the atmosphere and disperse over a wide area. These results can be applied to spacecraft launchpad accidents that expose spacecraft to such fires; during these fires, there is concern that some of the plutonium from the spacecraft power system will be carried with the aerosols. Accident-related lofting of this material would be the net result of many contributing processes that are currently being evaluated. To resolve the complexity of fire processes, a self-consistent model of the ground-level and upper-level parts of the plume was determined by merging ground-level optical measurements of the fire with lidar measurements of the aerosol plume at height during a series of SRP fire tests that simulated propellant fire accident scenarios. On the basis of the measurements and model results, the Johns Hopkins University Applied Physics Laboratory (JHU/APL) team was able to estimate the amount of aluminum oxide (alumina) lofted into the atmosphere above the fire. The quantification of this ratio is critical for a complete understanding of accident scenarios, because contaminants are transported through the plume. This paper provides an estimate for the mass of alumina lofted into the air. PMID:27409023

  8. Numerical simulation of a liquid propellant rocket motor

    NASA Astrophysics Data System (ADS)

    Salvador, Nicolas M. C.; Morales, Marcelo M.; Migueis, Carlos E. S. S.; Bastos-Netto, Demétrio

    2001-03-01

    This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems. This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for compressible flows only, was modified to work proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of references available in this area. Here an existing code for compressible flow was used and primitive variables, the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to permit the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and external flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theoretical results and with other codes already validated and well accepted by the CFD community.

  9. Experiments in thermosensitive cavitation of a cryogenic rocket propellant surrogate

    NASA Astrophysics Data System (ADS)

    Kelly, Sean Benjamin

    Cavitation is a phase-change phenomenon that may appear in practical devices, often leading to loss of performance and possible physical damage. Of particular interest is the presence of cavitation in rocket engine pumps as the cryogenic fluids cavitate in impellers and inducers. Unlike water, which has been studied exhaustively, cryogenic fluids undergo cavitation with significant thermal effect. Past attempts at analyzing this behavior in water have led to poor predictive capability due to the lack of data in the regime defined as thermosensitive cavitation. Fluids flowing near their thermodynamic critical point have a liquid-vapor density ratio that is orders of magnitude less than typical experimental fluids, so that the traditional equation-of-state and cavitation models do not apply. Thermal effects in cavitation have not been fully investigated due to experimental difficulties handling cryogenics. This work investigates the physical effects of thermosensitive cavitation in a model representative of a turbopump inducer in a modern rocket engine. This is achieved by utilizing a room-temperature testing fluid that exhibits a thermal effect equivalent to that experienced by cryogenic propellants. Unsteady surface pressures and high speed imaging collected over the span of thermophysical regimes ranging from thermosensitive to isothermal cavitation offer both quantitative and qualitative insight into the physical process of thermal cavitation. Physical and thermodynamic effects are isolated to identify the source of cavity conditions, oscillations and growth/collapse behavior. Planar laser imaging offers an instantaneous look inside the vapor cavity and at the behavior of the boundary between the two-phase region and freestream liquid. Nondimensional parameters are explored, with cavitation numbers, Reynolds Numbers, coefficient of pressure and nondimensional temperature in a broad range. Results in the form of cavitation regime maps, Strouhal Number of cavity

  10. Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

    NASA Technical Reports Server (NTRS)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

  11. Nonideal detonation and initiation behavior of a composite solid rocket propellant. [HMX/AP/Al

    SciTech Connect

    Dick, J.J.

    1981-01-01

    Shock initiation and detonation behavior of an HMX/AP/Al rocket propellant were studied for nonideal character. Low detonation velocities and unusual shock initiation behavior were observed. Failure to propagate steady detonation in cylinders of the propellant was also noted.

  12. Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

    NASA Astrophysics Data System (ADS)

    Gill, W.; Cruz-Cabrera, A. A.; Donaldson, A. B.; Lim, J.; Sivathanu, Y.; Bystrom, E.; Haug, A.; Sharp, L.; Surmick, D. M.

    2014-11-01

    Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

  13. [Progress in the protective medicine against [correction of aganist] rocket propellents].

    PubMed

    Hu, W X; Tan, C Y; Tan, S J; Jiang, J

    1999-12-01

    To review the progress in the major assignment, the organization and implementation of protection against liquid rocket propellent. The safety detection methods of the rocket [correction of rocked] propellent in the launching field were also discussed. Three steps of the sanitation and protection of the liquid propellent, the toxicity and the toxicology of hydrazine on central nervous system, blood circulatory system, assimilation system, respiratory system, immune system, liver, kidney, eye, skin and its hereditary toxicology were described. In addition, the clinical types of poisoning, the current principle and the common ways of the prevention and treatment of hydrazine and nitrogen oxides poisoning were summarized. PMID:12434814

  14. Development of an advanced rocket propellant handler's suit.

    PubMed

    Doerr, D F

    2001-01-01

    Most launch vehicles and satellites in the US inventory rely upon the use of hypergolic rocket propellants, many of which are toxic to humans. These fuels and oxidizers, such as hydrazine and nitrogen tetroxide have threshold limit values as low as 0.01 PPM. It is essential to provide space workers handling these agents whole body protection as they are universally hazardous not only to the respiratory system, but the skin as well. This paper describes a new method for powering a whole body protective garment to assure the safety of ground servicing crews. A new technology has been developed through the small business innovative research program at the Kennedy Space Center. Currently, liquid air is used in the environmental control unit (ECU) that powers the propellant handlers suit (PHE). However, liquid air exhibits problems with attitude dependence, oxygen enrichment, and difficulty with reliable quantity measurement. The new technology employs the storage of the supply air as a supercritical gas. This method of air storage overcomes all of three problems above while maintaining high density storage at relatively low vessel pressures (<7000 kPa or approximately 1000 psi). A one hour prototype ECU was developed and tested to prove the feasibility of this concept. This was upgraded by the design of a larger supercritical dewar capable of holding 7 Kg of air, a supply which provides a 2 hour duration to the PHE. A third version is being developed to test the feasibility of replacing existing air cooling methodology with a liquid cooled garment for relief of heat stress in this warm Florida environment. Testing of the first one hour prototype yielded data comparable to the liquid air powered predecessor, but enjoyed advantages of attitude independence and oxygen level stability. Thermal data revealed heat stress relief at least as good as liquid air supplied units. The application of supercritical air technology to this whole body protective ensemble marked an

  15. Development of an advanced rocket propellant handler's suit

    NASA Astrophysics Data System (ADS)

    Doerr, DonaldF.

    2001-08-01

    Most launch vehicles and satellites in the US inventory rely upon the use of hypergolic rocket propellants, many of which are toxic to humans. These fuels and oxidizers, such as hydrazine and nitrogen tetroxide have threshold limit values as low as 0.01 PPM. It is essential to provide space workers handling these agents whole body protection as they are universally hazardous not only to the respiratory system, but the skin as well. This paper describes a new method for powering a whole body protective garment to assure the safety of ground servicing crews. A new technology has been developed through the small business innovative research program at the Kennedy Space Center. Currently, liquid air is used in the environmental control unit (ECU) that powers the propellant handlers suit (PHE). However, liquid air exhibits problems with attitude dependence, oxygen enrichment, and difficulty with reliable quantity measurement. The new technology employs the storage of the supply air as a supercritical gas. This method of air storage overcomes all of three problems above while maintaining high density storage at relatively low vessel pressures (<7000 kPa or ˜1000 psi). A one hour prototype ECU was developed and tested to prove the feasibility of this concept. This was upgraded by the design of a larger supercritical dewar capable of holding 7 Kg of air, a supply which provides a 2 hour duration to the PHE. A third version is being developed to test the feasibility of replacing existing air cooling methodology with a liquid cooled garment for relief of heat stress in this warm Florida environment. Testing of the first one hour prototype yielded data comprobable to the liquid air powered predecessor, but enjoyed advantages of attitude independence and oxygen level stability. Thermal data revealed heat stress relief at least as good as liquid air supplied units. The application of supercritical air technology to this whole body protective ensemble marked an advancement in

  16. 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.

  17. Characterization of booster-rocket propellants and their simulants

    SciTech Connect

    Weirick, L.J.

    1989-01-01

    A series of shock-loading experiments on a composite and an energietic propellant and there simulants was conducted on a light-gas gun. The initial objectives were to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall threshold at various impact velocities. The Hugoniot data measured for the propellants fit the Hugoniot curves provided by the manufacturer of the propellants extremely well and the Hugoniot curves developed for the simulants matched those of the propellants. Threshold pressures to initiate reactions in the composite and energetic propellants were found to be 40 and 3 kbars, respectively. In spall tests, the composite propellant and its simulant exhibited spall strengths around 0.25 and 0.18 kbar, respectively. The energetic propellant and its simulant were somewhat stronger with spall strengths just above 0.33 and 0.22 kbar. 12 refs., 6 figs., 6 tabs.

  18. Atmospheric Manmade Glowings Phenomena Observed During the Launches of Solid Propellant Rockets

    NASA Astrophysics Data System (ADS)

    Chernouss, S. A.; Platov, V. V.; Upspensky, M. V.; Alpatov, V. V.; Kirillov, A. S.

    2015-09-01

    Exotic types of luminosities observed in the upper atmosphere always take place during the launch and flight of solid-propellant rockets We consider a large-scale geometry and dynamic features of such phenomena also physics of the intense turquoise (blue-green) glow observed in twilight conditions in the region of missile flight. This study has been based on numerous observations of different rocket flights in the atmosphere over Russia and Scandinavia. Formation of the monoxide aluminum clouds observed in the upper atmosphere is a result of interaction of the exhausted propellant products with the atomic oxygen. The sunlight excited the monoxide aluminum EA1O*) resonance emissions in the atmosphere. Careful studies of spectra of the manmade luminosities during rocket launch/flight permit us to know chemical, thermal and mechanical processes in the atmosphere similar as it is doing in experiments with the artificial cloud release from sounding rockets in the high latitude atmosphere.

  19. Solid rocket propellant waste disposal/ingredient recovery study

    NASA Technical Reports Server (NTRS)

    Mcintosh, M. J.

    1976-01-01

    A comparison of facility and operating costs of alternate methods shows open burning to be the lowest cost incineration method of waste propellant disposal. The selection, development, and implementation of an acceptable alternate is recommended. The recovery of ingredients from waste propellant has the probability of being able to pay its way, and even show a profit, when large consistent quantities of composite propellant are available. Ingredients recovered from space shuttle waste propellant would be worth over $1.5 million. Open and controlled burning are both energy wasteful.

  20. Extension of a simplified computer program for analysis of solid-propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Sforzini, R. H.

    1973-01-01

    A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.

  1. A facility for testing the acoustic combustion instability characteristics of solid rocket propellants

    NASA Technical Reports Server (NTRS)

    Mathes, H. B.

    1980-01-01

    A facility is described that has been specifically designed for small-scale laboratory testing of solid rocket propellants. A description of the facility is provided which includes the general plan of the facility and features related to personnel safety. One of the major activities in the facility is testing solid rocket propellants for combustion response to acoustic perturbations. A detailed discussion of acoustic instability testing is given including specially designed combustion apparatus, data acquisition, and signal conditioning. Techniques of data reduction are reviewed and some of the instrumentation problems that arise in this type of testing are mentioned along with practical solutions.

  2. Design issues for lunar in situ aluminum/oxygen propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Meyer, Michael L.

    1992-01-01

    Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.

  3. Modeling of heat generation in ammonia-treated solid rocket propellant

    SciTech Connect

    Raun, R.L.; Isom, K.B.

    1995-06-01

    With the end of the Cold War, safe, environmentally sound separation, recycling, and disposal of ingredients in solid rocket propellants and munitions has become a national priority. One approach to demilitarize solid rocket propellants is treatment with ammonia. Ammonia extracts the oxidizers ammonium perchlorate and HMX, yielding a solid reside that is more suitable for incineration and less sensitive to impact and other modes of accidental initiation. Ammonia treatment of nitroglycerin-containing propellants is complicated by an exothermic reaction between ammonia and nitroglycerin. If not removed, the heat generated by this reaction can cause propellant ignition. To help design safe treatment processes, a model for the ammonia-propellant reaction was developed, which integrates transient energy and species conservation equations to simulate ammonia diffusion, heat generation, and heat flow in a propellant and in the solid residue resulting from ammonia treatment. It was calibrated using residue thickness and thermocouple data for one propellant. The calibrated model was used to predict conditions leading to ignition of thin propellant strips. The results agree well with experimental observations.

  4. Development of a miniature solid propellant rocket motor for use in plume simulation studies

    NASA Technical Reports Server (NTRS)

    Baran, W. J.

    1974-01-01

    A miniature solid propellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solid propellant motors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.

  5. Experimental research and design planning in the field of liquid-propellant rocket engines conducted between 1934 - 1944 by the followers of F. A. Tsander

    NASA Technical Reports Server (NTRS)

    Dushkin, L. S.

    1977-01-01

    The development of the following Liquid-Propellant Rocket Engines (LPRE) is reviewed: (1) an alcohol-oxygen single-firing LPRE for use in wingless and winged rockets, (2) a similar multifiring LPRE for use in rocket gliders, (3) a combined solid-liquid propellant rocket engine, and (4) an aircraft LPRE operating on nitric acid and kerosene.

  6. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    Thomas, Matt; Bossard, John; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

    2001-01-01

    This viewgraph presentation gives an overview of laser ignition technology for bipropellant rocket engines applications. The objectives of this project include: (1) the selection test chambers and flows; (2) definition of the laser ignition setup; (3) pulse format optimization; (4) fiber optic coupled laser ignition system analysis; and (5) chamber integration issues definition. The testing concludes that rocket combustion chamber laser ignition is imminent. Support technologies (multiplexing, window durability/cleaning, and fiber optic durability) are feasible.

  7. Development of high temperature materials for solid propellant rocket nozzle applications

    NASA Technical Reports Server (NTRS)

    Manning, C. R., Jr.; Lineback, L. D.

    1974-01-01

    Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

  8. Features of optical phenomena connected with launches of solid-propellant ballistic rockets

    NASA Astrophysics Data System (ADS)

    Platov, Yu. V.; Chernouss, S. A.; Alpatov, V. V.

    2013-04-01

    Specific optical phenomena observed in the upper atmosphere layers and connected with launches of powerful solid-propellant rockets are considered: the development of spherically symmetric gas-dust formations having the shape of an extending torus in the image plane and the formation of regions with intense blue-green (turquoise) glow observed under twilight conditions along a rocket's flight path. The development of clouds can be represented by the model of a strong explosion occurring at the stage separation of solid-propellant rockets in the upper atmosphere. A turquoise glow arises as a result of resonance scattering of solar radiation on AlO molecules that are formed when metallic aluminum in the composition of fuel interacts with atmosphere components and combustion products.

  9. Propellant acquisition for single-stage rocket technology

    NASA Astrophysics Data System (ADS)

    Grayson, G.; Distefano, E.

    1993-06-01

    An analytical design approach defines the liquid oxygen (LO2) propellant-acquisition system for the vertical-take-off and vertical-landing DC-X flight test vehicle. The DC-X trajectory includes a rotation maneuver in which the vehicle pitches from a nose-first to a tail-first orientation. The 8-ft diameter LO2 tank accommodates outflow rates as high as 120 lbm/sec, while vehicle drag forces result in a lateral acceleration approaching 25 ft/sq sec during rotation. FLOW-3D computational results show that, without propellant control, these conditions cause the LO2 to slosh within the tank and expose the outlet to ullage gas during the rotation maneuver. Using FLOW 3-D simulations of several baffle configurations, one propellant acquisition concept is selected. In this concept, a simple cone-shaped containment baffle compartmentalizes the tank and an outlet extension offsets the outlet to where the propellant is biased. The final design incorporates a factor of safety greater than 2 resulting in continuous coverage of the outlet by propellant as validated by FLOW-3D simulation.

  10. Lead-Free Double-Base Propellant for the 2.75 Inch Rocket Motor

    NASA Technical Reports Server (NTRS)

    Magill, B. T.; Nauflett, G. W.; Furrow, K. W.

    2000-01-01

    The current MK 66 2.75 inch Rocket Motor double-base propellant contains the lead-based ballistic modifier LC-12-15 to achieve the desired plateau and mesa burning rate characteristics. The use of lead compounds poses a concern for the environment and for personal safety due to the metal's toxic nature when introduced into the atmosphere by propellant manufacture, rocket motor firing, and disposal. Copper beta-resorcylate (copper 2,4-di-hydroxy-benzoate) was successfully used in propellant as a simple modifier in the mid 1970's. This and other compounds have also been mixed with lead salts to obtain more beneficial ballistic results. Synthesized complexes of lead and copper compounds soon replaced the mixtures. The complexes incorporate the lead, copper lack of organic liquids, which allows for easier propellant processing. About ten years ago, the Indian Head Division, Naval Surface Warfare Center (NSWC), initiated an effort to develop a lead-free propellant for use in missile systems. Several lead-free propellant candidate formulations were developed. About five years ago, NSWC, in conjunction with Alliant Techsystems, Radford Army Ammunition Plant, continued ballistic modifier investigations. A four component ballistic modifier system without lead for double-base propellants that provide adequate plateau and mesa burn rate characteristics was developed and patented. The ballistic modifier's system contains bismuth subsalicylate, 1.5 percent; copper salicylate, 1.0 percent, copper stannate, 0.77 percent; and carbon black, 0.1 percent. Action time and impulse data obtained through multiple static firings indicate that the new lead-free double-base propellant, while not a match for NOSIH-AA-2, will be a very suitable replacement in the 2.75 inch Rocket Motor. Accelerated aging of the double-base propellant containing the lead-free ballistic modifier showed that it had a much higher rate of stabilizer depletion than the AA-2. A comprehensive study showed that an

  11. 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.

  12. Molded composite pyrogen igniter for rocket motors. [solid propellant ignition

    NASA Technical Reports Server (NTRS)

    Heier, W. C.; Lucy, M. H. (Inventor)

    1978-01-01

    A lightweight pyrogen igniter assembly including an elongated molded plastic tube adapted to contain a pyrogen charge was designed for insertion into a rocket motor casing for ignition of the rocket motor charge. A molded plastic closure cap provided for the elongated tube includes an ignition charge for igniting the pyrogen charge and an electrically actuated ignition squib for igniting the ignition charge. The ignition charge is contained within a portion of the closure cap, and it is retained therein by a noncorrosive ignition pellet retainer or screen which is adapted to rest on a shoulder of the elongated tube when the closure cap and tube are assembled together. A circumferentially disposed metal ring is provided along the external circumference of the closure cap and is molded or captured within the plastic cap in the molding process to provide, along with O-ring seals, a leakproof rotary joint.

  13. Rocket propellant inhalation in the Apollo-Soyuz astronauts.

    PubMed

    DeJournette, R L

    1977-10-01

    Acute exposure to monomethylhydrazine and dinitrogen tetroxide, the principal toxic irritants in rocket fuels, is described with particular attention to the development of pulmonary edema as a herbinger of more severe central nervous system toxicity. An acute respiratory embarrassment is documented and possible means of therapy based on animal experimental models is suggested. Early clinical and radiographic examination as a baseline for further evaluation is essential, with follow-up radiographs recommended for assessment of possible developing chronic lung disease. PMID:897171

  14. 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.

  15. System for imposing directional stability on a rocket-propelled vehicle

    NASA Technical Reports Server (NTRS)

    Perkins, H. (Inventor)

    1976-01-01

    An improved system for use in imposing directional stability on a rocket-propelled vehicle is described. The system includes a pivotally supported engine-mounting platform, a gimbal ring mounted on the platform and adapted to pivotally support a rocket engine and an hydraulic actuator connected to the platform for imparting selected pivotal motion. An accelerometer and a signal comparator circuit for providing error intelligence indicative of aberration in vehicle acceleration is included along with an actuator control circuit connected with the actuator and responsive to error intelligence for imparting pivotal motion to the platform. Relocation of the engine's thrust vector is thus achieved for imparting directional stability to the vehicle.

  16. Transferring jet engine diagnostic and control technology to liquid propellant rocket engines

    SciTech Connect

    Alcock, J.F.; Hagar, S.K.

    1989-01-01

    This paper presents the methodology for developing a diagnostic and control system for a current, operational jet engine. A description is given of each development stage, the system components and the technologies which could be transferred to liquid propellant rocket engines. Finally, the operational impact is described in terms of cost and maintenance based on actual jet engine experience. Efforts are continuing to develop new diagnostic techniques under IR D for application on the advanced technical fighter. Already improved techniques and application methods are becoming available. This technology is being evaluated and may also be transferred to rocket engine diagnostic and control system development.

  17. Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

    NASA Technical Reports Server (NTRS)

    Hulka, James R.; Jones, Gregg W.

    2010-01-01

    Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using

  18. Acceleration effects on the performance of solid-propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Lucy, M. H.; Jones, I. W.; Stephens, M. V.

    1976-01-01

    Some acceleration effects on rocket performance have been well publicized. The dynamic process, characterized by marked increases in 'localized' burning rate, produces excessive case heating, slag retention, pressure buildup, and/or internal flow alterations. Data are presented illustrating drastic effects at low accelerations for sustainer type propellants and its relevance to several recent failures. Normalized orientation dependence of rate augmentation appears coupled to acceleration level and base burning rate. Effects appear influenced by propellant composition. Predictions using subscale motor data show good agreement with observed performance for ground spin and flight tests. Subscale test methods and results are also discussed.

  19. Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Meacham, Michael B.; Kennett, Andrew; Townsend, Derik J.; Marti, Benjamin

    2013-01-01

    Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

  20. Computational analysis of liquid hypergolic propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Krishnan, A.; Przekwas, A. J.; Gross, K. W.

    1992-01-01

    The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

  1. A study on various methods of supplying propellant to an orbit insertion rocket engine

    NASA Technical Reports Server (NTRS)

    Boretz, J. E.; Huniu, S.; Thompson, M.; Pagani, M.; Paulsen, B.; Lewis, J.; Paul, D.

    1980-01-01

    Various types of pumps and pump drives were evaluated to determine the lightest weight system for supplying propellants to a planetary orbit insertion rocket engine. From these analyses four candidate propellant feed systems were identified. Systems Nos. 1 and 2 were both battery powered (lithium-thionyl-chloride or silver-zinc) motor driven pumps. System 3 was a monopropellant gas generator powered turbopump. System 4 was a bipropellant gas generator powered turbopump. Parameters considered were pump break horsepower, weight, reliability, transient response and system stability. Figures of merit were established and the ranking of the candidate systems was determined. Conceptual designs were prepared for typical motor driven pumps and turbopump configurations for a 1000 lbf thrust rocket engine.

  2. Biogenic technology for recultivation of lands contaminated due to rocket propellant spillage

    NASA Astrophysics Data System (ADS)

    Kovshov, S. V.; Garkushev, A. U.; Sazykin, A. M.

    2015-04-01

    This article describes the problem of soil properties deterioration due to rocket propellant spillage. Melange and samin are considered to be the main pollutants. Provision is made for assessment of the existing mechanisms for monitoring of quality and recultivation of lands disturbed by rocket propellant spills. Some major disadvantages of currently used standard recultivation technologies are listed. An alternative is the use of more environmentally safe and cost effective methods aimed at disturbed lands biological restoration. An example of such a technology is covering the affected area with a biogenic mixture consisting of biohumus and sodium carboxymethyl cellulose followed by seeding it with specially selected herbal mixtures. It was found out that the most rational parameters of such protective layer is its thickness of 3 cm, and 99:1 ratio of its constituent components.

  3. Prediction of explosive yield and other characteristics of liquid rocket propellant explosions

    NASA Technical Reports Server (NTRS)

    Farber, E. A.; Smith, J. H.; Watts, E. H.

    1973-01-01

    Work which has been done at the University of Florida in arriving at credible explosive yield values for liquid rocket propellants is presented. The results are based upon logical methods which have been well worked out theoretically and verified through experimental procedures. Three independent methods to predict explosive yield values for liquid rocket propellants are described. All three give the same end result even though they utilize different parameters and procedures. They are: (1) mathematical model; (2) seven chart approach; and (3) critical mass method. A brief description of the methods, how they were derived, how they were applied, and the results which they produced are given. The experimental work used to support and verify the above methods both in the laboratory and in the field with actually explosive mixtures are presented. The methods developed are used and their value demonstrated in analyzing real problems, among them the destruct system of the Saturn 5, and the early configurations of the space shuttle.

  4. Regeneratively cooled rocket engine for space storable propellants

    NASA Technical Reports Server (NTRS)

    Wagner, W. R.

    1973-01-01

    Analysis, design, fabrication, and test efforts were performed for the existing OF2/B2H6 regeneratively cooled lK (4448 N) thrust chamber to illustrate simultaneous B2H6 fuel and OF2 oxidizer cooling and to provide results for a gaseous propellant condition injected into the combustion chamber. Data derived from performance, thermal and flow measurements confirmed predictions derived from previous test work and from concurrent analytical study. Development data derived from the experimental study were indicated to be sufficient to develop a preflight thrust chamber demonstrator prototype for future space mission objectives.

  5. Regeneratively cooled rocket engine for space storable propellants

    NASA Technical Reports Server (NTRS)

    Wagner, W. R.; Waldman, B. J.

    1973-01-01

    Analyses and experimental studies were performed with the OF2 (F2/O2)/B2H6 propellant combination over a range in operating conditions to determine suitability for a space storable pressure fed engine configuration for an extended flight space vehicle configuration. The regenerative cooling mode selected for the thrust chamber was explored in detail with the use of both the fuel and oxidizer as coolants in an advanced milled channel construction thrust chamber design operating at 100 psia chamber pressure and a nominal mixture ratio of 3.0 with a 60:1 area ratio nozzle. Benefits of the simultaneous cooling as related to gaseous injection of both fuel and oxidizer propellants were defined. Heat transfer rates, performance and combustor stability were developed for impinging element triplet injectors in uncooled copper calorimeter hardware with flow, pressure and temperature instrumentation. Evaluation of the capabilities of the B2H6 and OF2 during analytical studies and numerous tests with flow through electrically heated blocks provided design criteria for subsequent regenerative chamber design and fabrication.

  6. Attenuation studies of booster-rocket propellants and their simulants

    SciTech Connect

    Weirick, L.J.

    1990-08-01

    A series of impact experiments on a composite propellant, an energetic propellant, and their simulants was recently completed using a light-gas gun. Previous experiments were done to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall damage at various impact velocities. The present studies measured the attenuation of shock waves in these materials, completing the shock characterization needed for material modeling. An initial impulse of 2.0 GPa magnitude and {approximately}0.6 {mu}s duration was imposed upon samples of various thicknesses. VISAR was used to measure the free-surface velocity at the back of the samples; these data were used to generate a curve of shock-wave attenuation versus sample thickness for each material. Results showed that all four materials attenuated the shock wave very similarly. Material thicknesses of 3.0, 7.62, 12.7, and 19.0 mm attenuated the shock wave {approximately}16%, 33%, 50%, and 66% respectively. 14 refs., 12 figs., 4 tabs.

  7. Measured particulate behavior in a subscale solid propellant rocket motor

    NASA Astrophysics Data System (ADS)

    Brennan, W. D.; Hovland, D. L.; Netzer, D. W.

    1992-10-01

    Particulate matter are sized in the exhaust nozzle and plume of small rocket motors of varying geometry to assess the effects of the expansion process on particle size. Both converging and converging-diverging nozzles are considered, and particle sizing is accomplished at pressures of up to 4.36 MPa with aluminum loadings of 2.0 and 4.7 percent. An instrument based on Fraunhofer diffraction is used to measure the particle-size distributions showing that: (1) high burning rates reduce particle agglomeration and increase C* efficiency; (2) high pressures lead to small and monomodal D32 entering the nozzle; and (3) D32 sizes increase appreciably at the tailoff. Some variations in plume signature are theorized to be caused by the tailoff phenomenon, and particle collisions and/or surface effects in the nozzle convergence are suggested by the reduced number of larger particles at the nozzle convergence.

  8. Low loss injector for liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Vonpragenau, G. L. (Inventor)

    1986-01-01

    A low pressure loss injector element is disclosed for the main combustion chamber of a rocket engine which includes a lox post terminating in a cylindrical barrel. Received within the barrel is a lox plug which is threaded in the lox post and includes an interchangeable lox metering sieve which meters the lox into an annular lox passage. A second annular gas passage is coaxial with the annular lox passage. A cylindrical sleeve surrounds the annular gas passage and includes an interchangeable gas metering seive having metering orifices through which a hot gas passes into the annular passage. The jets which emerge from the annular lox passage and annular gas passage intersect in a recessed area away from the combustion area. Thus, mixing and combustion stability are enhanced.

  9. Bleed cycle propellant pumping in a gas-core nuclear rocket engine system

    NASA Technical Reports Server (NTRS)

    Kascak, A. F.; Easley, A. J.

    1972-01-01

    The performance of ideal and real staged primary propellant pumps and bleed-powered turbines was calculated for gas-core nuclear rocket engines over a range of operating pressures from 500 to 5000 atm. This study showed that for a required engine operating pressure of 1000 atm the pump work was about 0.8 hp/(lb/sec), the specific impulse penalty resulting from the turbine propellant bleed flow as about 10 percent; and the heat required to preheat the propellant was about 7.8 MN/(lb/sec). For a specific impulse above 2400 sec, there is an excess of energy available in the moderator due to the gamma and neutron heating that occurs there. Possible alternative pumping cycles are the Rankine or Brayton cycles.

  10. Alternate propellants for the space shuttle solid rocket booster motors. [for reducing environmental impact of launches

    NASA Technical Reports Server (NTRS)

    1973-01-01

    As part of the Shuttle Exhaust Effects Panel (SEEP) program for fiscal year 1973, a limited study was performed to determine the feasibility of minimizing the environmental impact associated with the operation of the solid rocket booster motors (SRBMs) in projected space shuttle launches. Eleven hypothetical and two existing limited-experience propellants were evaluated as possible alternates to a well-proven state-of-the-art reference propellant with respect to reducing emissions of primary concern: namely, hydrogen chloride (HCl) and aluminum oxide (Al2O3). The study showed that it would be possible to develop a new propellant to effect a considerable reduction of HCl or Al2O3 emissions. At the one extreme, a 23% reduction of HCl is possible along with a ll% reduction in Al2O3, whereas, at the other extreme, a 75% reduction of Al2O3 is possible, but with a resultant 5% increase in HCl.

  11. Shock-wave characterization of energetic booster-rocket propellant WAK-2 and its simulant UGS

    SciTech Connect

    Weirick, L.J.

    1989-10-01

    A series of shock-loading experiments on an energetic propellant and its simulant was conducted on a light-gas gun. The purpose of this work was to characterize the shock sensitivity of WAK-2, which is a composite-modified, double-based, booster-rocket propellant and its simulant UGS. The initial objectives were to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall threshold at various impact velocities. The Hugoniot data obtained for the propellant fits the Hugoniot curve provided by the manufacturer of the propellant. A Hugoniot curve developed for the simulant was found to match that of the propellant. The initial density, {rho}{sub 0}, initial bulk sound velocity, C{sub 0}, and constant S values for the energetic propellant WAK-2 and its simulant UGS were 1.85 g/cm{sup 3}, 2.2 mm/{mu}s and 2.66, respectively. The ignition threshold pressure of the WAK-2 was found to be in the range of 3 kbar. A violent reaction was observed for a sample impacted at a pressure of 22 kbar. In spall tests, impact pressures in the range of 1.1 to 3.1 kbar were applied to the propellant/simulant. The propellant exhibited spall strengths {approximately}0.33 kbar, with its simulant being somewhat weaker, {approximately}0.22 kbar. Scanning electron microscopy and electron microprobe analysis were used to characterize the microstructures of the materials and to determine the details of the spall events. 20 refs., 19 figs., 9 tabs.

  12. Studies of Fission Fragment Rocket Engine Propelled Spacecraft

    NASA Technical Reports Server (NTRS)

    Werka, Robert O.; Clark, Rodney; Sheldon, Rob; Percy, Thomas K.

    2014-01-01

    The NASA Office of Chief Technologist has funded from FY11 through FY14 successive studies of the physics, design, and spacecraft integration of a Fission Fragment Rocket Engine (FFRE) that directly converts the momentum of fission fragments continuously into spacecraft momentum at a theoretical specific impulse above one million seconds. While others have promised future propulsion advances if only you have the patience, the FFRE requires no waiting, no advances in physics and no advances in manufacturing processes. Such an engine unequivocally can create a new era of space exploration that can change spacecraft operation. The NIAC (NASA Institute for Advanced Concepts) Program Phase 1 study of FY11 first investigated how the revolutionary FFRE technology could be integrated into an advanced spacecraft. The FFRE combines existent technologies of low density fissioning dust trapped electrostatically and high field strength superconducting magnets for beam management. By organizing the nuclear core material to permit sufficient mean free path for escape of the fission fragments and by collimating the beam, this study showed the FFRE could convert nuclear power to thrust directly and efficiently at a delivered specific impulse of 527,000 seconds. The FY13 study showed that, without increasing the reactor power, adding a neutral gas to the fission fragment beam significantly increased the FFRE thrust through in a manner analogous to a jet engine afterburner. This frictional interaction of gas and beam resulted in an engine that continuously produced 1000 pound force of thrust at a delivered impulse of 32,000 seconds, thereby reducing the currently studied DRM 5 round trip mission to Mars from 3 years to 260 days. By decreasing the gas addition, this same engine can be tailored for much lower thrust at much higher impulse to match missions to more distant destinations. These studies created host spacecraft concepts configured for manned round trip journeys. While the

  13. New Frontiers AO: Advanced Materials Bi-propellant Rocket (AMBR) Engine Information Summary

    NASA Technical Reports Server (NTRS)

    Liou, Larry C.

    2008-01-01

    The Advanced Material Bi-propellant Rocket (AMBR) engine is a high performance (I(sub sp)), higher thrust, radiation cooled, storable bi-propellant space engine of the same physical envelope as the High Performance Apogee Thruster (HiPAT(TradeMark)). To provide further information about the AMBR engine, this document provides details on performance, development, mission implementation, key spacecraft integration considerations, project participants and approach, contact information, system specifications, and a list of references. The In-Space Propulsion Technology (ISPT) project team at NASA Glenn Research Center (GRC) leads the technology development of the AMBR engine. Their NASA partners were Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL). Aerojet leads the industrial partners selected competitively for the technology development via the NASA Research Announcement (NRA) process.

  14. Test data from small solid propellant rocket motor plume measurements (FA-21)

    NASA Technical Reports Server (NTRS)

    Hair, L. M.; Somers, R. E.

    1976-01-01

    A program is described for obtaining a reliable, parametric set of measurements in the exhaust plumes of solid propellant rocket motors. Plume measurements included pressures, temperatures, forces, heat transfer rates, particle sampling, and high-speed movies. Approximately 210,000 digital data points and 15,000 movie frames were acquired. Measurements were made at points in the plumes via rake-mounted probes, and on the surface of a large plate impinged by the exhaust plume. Parametric variations were made in pressure altitude, propellant aluminum loading, impinged plate incidence angle and distance from nozzle exit to plate or rake. Reliability was incorporated by continual use of repeat runs. The test setup of the various hardware items is described along with an account of test procedures. Test results and data accuracy are discussed. Format of the data presentation is detailed. Complete data are included in the appendix.

  15. Nonlinear analysis of longitudinal-mode liquid propellant rocket combustion instability

    NASA Technical Reports Server (NTRS)

    Jeng, San-Mou; Litchford, Ron J.

    1991-01-01

    A computational technique for the nonlinear analysis of longitudinal-mode liquid propellant rocket combustion instability is examined based on the unsteady, quasi-one-dimensional Euler equations with appropriate source terms introduced to account for interphase transport coupling with the spray. The method is first assessed for unsteady, nonreacting, isentropic duct flow with specified admittances at the outflow and inflow boundaries. For small amplitude disturbances, numerical results based on a sufficient number of grid points compare favorably with predictions from a small-disturbance linearized analysis.

  16. Solid propellant rocket motor internal ballistics performance variation analysis, phase 3

    NASA Technical Reports Server (NTRS)

    Sforzini, R. H.; Foster, W. A., Jr.; Murph, J. E.; Adams, G. W., Jr.

    1977-01-01

    Results of research aimed at improving the predictability of off nominal internal ballistics performance of solid propellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.

  17. Effect of silicone oil on nozzle heat transfer in solid propellant rockets

    NASA Technical Reports Server (NTRS)

    Ramohalli, K.

    1978-01-01

    The feasibility of reducing heat transfer rates to inert nozzle walls in solid rockets by addition of silicone oil as a propellant ingredient is explored experimentally. A metallized control propellant and its counterpart with 1% silicone oil are fired in a 2 1/2-in. diameter end burning all-metal burner. Pressure-time histories are obtained; temperature-time data at the nozzle throat are obtained in some of the runs. The heat loss by the product gases to the cold inert walls of the nozzle lower the temperature of the gases leading to the condensation of aluminum oxide on the walls. This condensation is used as an index of the heat transfer rate to the nozzle. On a statistical basis it is found that the silicone propellant gives practically no condensation in 70% of the firings contrasted with 21% for its counterpart with no silicone oil in it. The gas temperature at the nozzle throat is experimentally seen to be generally higher with the silicone propellant than with its counterpart; this higher temperature is consistent with a preliminary analysis that is used here to relate the nozzle surface temperature to the heat transfer from the product gases.

  18. Effect of silicone oil on solid propellant combustion in small motors. [for rockets

    NASA Technical Reports Server (NTRS)

    Ramohalli, K.

    1980-01-01

    The feasibility of reducing troublesome nozzle blockage (by condensation deposits) in laboratory-scale solid rockets by addition of a silicone oil as a propellant ingredient was explored experimentally. An aluminized composite propellant and its counterpart with 1% silicone oil replacing part of the binder were fired in a 63.5 mm diameter, end-burning, all-metal burner. Pressure-time histories were recorded for all of the tests by a Taber gauge mounted at the downstream end of the chamber; temperature-time data at the nozzle throat were obtained in some of the runs by thermocouples having junctions positioned at the wall but insulated from the metal. Deposition of condensables on the nozzle walls causing a progressive increase in the chamber pressure with time was noted. The fraction of firings exhibiting practically no condensation was 59% with silicone and 32% without. On the average, temperature readings at the nozzle throat were higher with the silicone propellants. Although various phenomena may contribute to these findings, the results are not understood completely.

  19. Characterization of rocket propellant combustion products: Description of sampling and analysis methods for rocket exhaust characterization studies

    SciTech Connect

    Jenkins, R.A.

    1990-06-07

    A systematic approach has been developed and experimentally validated for the sampling and chemical characterization of the rocket motor exhaust generated from the firing of scaled down test motors at the US Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama. The overall strategy was to sample and analyze major exhaust constituents in near real time, while performing off-site analyses of samples collected for the determination of trace constituents of the particulate and vapor phases. Initial interference studies were performed using atmospheric pressure burns of 1 g quantities of propellants in small chambers at Oak Ridge National Laboratory. Carbon monoxide and carbon dioxide were determined using non-dispersive infrared instrumentation. Hydrogen cyanide, hydrogen chloride, and ammonia determinations were made using ion selective electrode technology. Oxides of nitrogen were determined using chemiluminescence instrumentation. Airborne particulate mass concentration was determined using infrared forward scattering measurements and a tapered element oscillating microbalance, as well as conventional gravimetry. Particulate phase metals were determined by collection on Teflon membrane filters, followed by inductively coupled plasma and atomic absorption analysis. Particulate phase polynuclear aromatic hydrocarbons (PAH) and nitro-PAH were collected using high volume sampling on a two stage filter. Target species were extracted, and quantified by gas chromatography/mass spectrometry (GC/MS). Vapor phase species were collected on multi-sorbent resin traps, and subjected to thermal desorption GC/MS for analysis. 11 refs., 1 fig., 1 tab.

  20. A Preliminary Investigation on the Destruction of Solid-Propellant Rocket Motors by Impact from Small Particles

    NASA Technical Reports Server (NTRS)

    Carter, David J., Jr.

    1960-01-01

    An investigation was conducted to determine whether solid-propellant rocket motors could be ignited and destroyed by small-particle impacts at particle velocities up to a approximately 10,940 feet per second. Spheres ranging from 1/16 to 7/32 inch in diameter were fired into simulated rocket motors containing T-22 propellant over a range of ambient pressures from sea level to 0.12 inch of mercury absolute. Simulated cases of stainless steel, aluminum alloy, and laminated Fiberglas varied in thickness from 1/50 to 1/8 inch. Within the scope of this investigation, it was found that ignition and explosive destruction of simulated steel-case rocket motors could result from impacts by steel spheres at the lowest attainable pressure.

  1. Inflatable mandrel fabrication technology - Advantages for the containment of rocket propellants

    NASA Astrophysics Data System (ADS)

    Moser, Daniel J.

    1992-07-01

    This paper discusses and compares the attributes of various mandrel types as they pertain to the fabrication of filament-wound composite containment vessels for rocket propellants, solid or liquid. The issues of dimensional conformity, processing parameters, unit costs, vessel performance, and development lead times are raised. The continuous fiber reinforced, reusable inflatable mandrel concept is explained and shown to have unique advantages over more traditional mandrel types. Burst pressure performance is equivalent to, or slightly better than, the results from vessels built on the more conventional net metal mandrel. The co-cured insulator process used in conjunction with the inflatable mandrel is shown to be superior in some respects. Some experimental findings are presented along with description of the processing parameters that must be understood when using inflatable mandrels.

  2. Theoretical Performance of Liquid Hydrogen with Liquid Oxygen as a Rocket Propellant

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; McBride, Bonnie J.

    1959-01-01

    Theoretical rocket performance for both equilibrium and frozen composition during expansion was calculated for the propellant combination liquid hydrogen and liquid oxygen at four chamber pressures (60, 150, 300, and 600 lb/sq in. abs) and a wide range of pressure ratios (1 to 4000) and oxidant-fuel ratios (1.190 to 39.683). Data are given to estimate performance parameters at chamber pressures other than those for which data are tabulated. The parameters included are specific impulse, specific impulse in vacuum, combustion-chamber temperature, nozzle-exit temperature, molecular weight, molecular-weight derivatives, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, Mach number, and equilibrium gas compositions.

  3. Theoretical performance of liquid ammonia and liquid fluorine as a rocket propellant

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Huff, Vearl N

    1953-01-01

    Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 311.5 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

  4. Inverse synthetic aperture radar imagery of a man with a rocket propelled grenade launcher

    NASA Astrophysics Data System (ADS)

    Tran, Chi N.; Innocenti, Roberto; Kirose, Getachew; Ranney, Kenneth I.; Smith, Gregory

    2004-08-01

    As the Army moves toward more lightly armored Future Combat System (FCS) vehicles, enemy personnel will present an increasing threat to U.S. soldiers. In particular, they face a very real threat from adversaries using shoulder-launched, rocket propelled grenade (RPG). The Army Research Laboratory has utilized its Aberdeen Proving Ground (APG) turntable facility to collect very high resolution, fully polarimetric Ka band radar data at low depression angles of a man holding an RPG. In this paper, we examine the resulting low resolution and high resolution range profiles; and based on the observed radar cross section (RCS) value, we attempt to determine the utility of Ka band radar for detecting enemy personnel carrying RPG launchers.

  5. Characterization of typical platelet injector flow configurations. [liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Hickox, C. E.

    1975-01-01

    A study to investigate the hydraulic atomization characteristics of several novel injector designs for use in liquid propellant rocket engines is presented. The injectors were manufactured from a series of thin stainless steel platelets through which orifices were very accurately formed by a photoetching process. These individual platelets were stacked together and the orifices aligned so as to produce flow passages of prescribed geometry. After alignment, the platelets were bonded into a single, 'platelet injector', unit by a diffusion bonding process. Because of the complex nature of the flow associated with platelet injectors, it was necessary to use experimental techniques, exclusively, throughout the study. Large scale models of the injectors were constructed from aluminum plates and the appropriate fluids were modeled using a glycerol-water solution. Stop-action photographs of test configurations, using spark-shadowgraph or stroboscopic back-lighting, are shown.

  6. Theoretical performance of liquid hydrogen and liquid fluorine as a rocket propellant

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Huff, Vearl N

    1953-01-01

    Theoretical values of performance parameters for liquid hydrogen and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ration of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 364.6 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

  7. Solid rocket booster performance evaluation model. Volume 3: Sample case. [propellant combustion simulation/internal ballistics

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The solid rocket booster performance evaluation model (SRB-11) is used to predict internal ballistics in a sample motor. This motor contains a five segmented grain. The first segment has a 14 pointed star configuration with a web which wraps partially around the forward dome. The other segments are circular in cross-section and are tapered along the interior burning surface. Two of the segments are inhibited on the forward face. The nozzle is not assumed to be submerged. The performance prediction is broken into two simulation parts: the delivered end item specific impulse and the propellant properties which are required as inputs for the internal ballistics module are determined; and the internal ballistics for the entire burn duration of the motor are simulated.

  8. 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.}

  9. 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.

  10. Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

    NASA Technical Reports Server (NTRS)

    Kenny, Jeremy R.; Moser, Marlow D.; Hulka, James; Jones, Gregg

    2006-01-01

    Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with fixed designs and is a critical requirement in lunar and other planetary landing missions. A task in the Constellation University Institutes Program (CUIP) has been designed to evaluate spray characteristics when liquid propellant rocket engine injectors are scaled and throttled. The specific objectives of the present study are to characterize injection and primary atomization using cold flow simulations of the reacting sprays. These simulations can provide relevant information because the injection and primary atomization are believed to be the spray processes least affected by the propellant reaction. Cold flow studies also provide acceptable test conditions for a university environment. Three geometric scales - 1/4- scale, 1/2-scale, and full-scale - of two different injector element types - swirl coaxial and shear coaxial - will be designed, fabricated, and tested. A literature review is currently being conducted to revisit and compile the previous scaling documentation. Because it is simple to perform, throttling will also be examined in the present work by measuring primary atomization characteristics as the mass flow rate and pressure drop of the six injector element concepts are reduced, with corresponding changes in chamber backpressure. Simulants will include water and gaseous nitrogen, and an optically accessible chamber will be used for visual and laser-based diagnostics. The chamber will include curtain flow capability to repress recirculation, and additional gas injection to provide independent control of the

  11. 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.

  12. Amplification of Reynolds number dependent processes by wave distortion. [acoustic instability of liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Ventrice, M. B.; Fang, J. C.; Purdy, K. R.

    1975-01-01

    A system using a hot-wire transducer as an analog of a liquid droplet of propellant was employed to investigate the ingredients of the acoustic instability of liquid-propellant rocket engines. It was assumed that the combustion process was vaporization-limited and that the combustion chamber was acoustically similar to a closed-closed right-circular cylinder. Before studying the hot-wire closed-loop system (the analog system), a microphone closed-loop system, which used the response of a microphone as the source of a linear feedback exciting signal, was investigated to establish the characteristics of self-sustenance of acoustic fields. Self-sustained acoustic fields were found to occur only at resonant frequencies of the chamber. In the hot-wire closed-loop system, the response of hot-wire anemometer was used as the source of the feedback exciting signal. The self-sustained acoustic fields which developed in the system were always found to be harmonically distorted and to have as their fundamental frquency a resonant frequency for which there also existed a second resonant frequency which was approximately twice the fundamental frequency.

  13. Solid-propellant rocket motor internal ballistics performance variation analysis, phase 5

    NASA Technical Reports Server (NTRS)

    Sforzini, R. H.; Murph, J. E.

    1980-01-01

    The results of research aimed at improving the predictability of internal ballistics performance of solid-propellant rocket motors (SRM's) including thrust imbalance between two SRM's firing in parallel are presented. Static test data from the first six Space Shuttle SRM's is analyzed using a computer program previously developed for this purpose. The program permits intentional minor design biases affecting the imbalance between any two SMR's to be removed. Results for the last four of the six SRM's, with only the propellant bulk temperature as a non-random variable, are generally within limits predicted by theory. Extended studies of internal ballistic performance of single SRM's are presented based on an earlier developed mathematical model which includes an assessment of grain deformation. The erosive burning rate law used in the model is upgraded and made more general. Excellent results are obtained in predictions of the performances of five different SRM's of quite different sizes and configurations. These SRM's all employ PBAN type propellants with ammonium perchlorate oxidizer and 16 to 20% aluminum except one which uses carboxyl terminated butadiene binder. The only non-calculated parameters in the burning rate equations that are changed for the different SRM's are the zero crossflow velocity burning rate coefficients and exponents. The results, in general, confirm the importance of grain deformation. The improved internal ballistic model makes practical development of an effective computer program for application of an optimization technique to SRM design which is also demonstrated. The program uses a pattern search technique to minimize the difference between a desired thrust-time trace and one calculated based on the internal ballistic model.

  14. 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

  15. Flight Investigation of the Performance of a Two-stage Solid-propellant Nike-deacon (DAN) Meteorological Sounding Rocket

    NASA Technical Reports Server (NTRS)

    Heitkotter, Robert H

    1956-01-01

    A flight investigation of two Nike-Deacon (DAN) two-stage solid-propellant rocket vehicles indicated satisfactory performance may be expected from the DAN meteorological sounding rocket. Peak altitudes of 356,000 and 350,000 feet, respectively, were recorded for the two flight tests when both vehicles were launched from sea level at an elevation angle of 75 degrees. Performance calculations based on flight-test results show that altitudes between 358,000 feet and 487,000 feet may be attained with payloads varying between 60 pounds and 10 pounds.

  16. 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.

  17. Rotational inviscid flow in laterally burning solid-propellant rocket motors

    NASA Astrophysics Data System (ADS)

    Balakrishnan, G.; Linan, A.; Williams, F. A.

    1992-12-01

    A theoretical analysis to determine the effects of mass addition on the inviscid but rotational and compressible flowfield in a porous duct with the injection rate dependent on the local pressure is performed for large ratios of length-to-duct diameter. The problem of describing the flow is reduced to the solution of a single integral equation. The ratio of specific heat gamma, and a constant pressure exponent n, measuring the dependence of the rate of mass injection on the local pressure, are the parameters of the solutions. The integral equation is solved numerically, and parametric results are presented for gamma, varying from 1 to 5/3 and for n varying from 0 to 1. A choking phenomenon is exhibited at a critical length of the duct in the vicinity of which the Mach number approaches unity. The choking condition, which is relevant to the operation of nozzleless solid-propellant rocket motors, is obtained parametrically in the present study and compared with corresponding results for irrotational, quasi-one-dimensional flow. The rotationality reduces the choking pressure.

  18. NASTRAN cyclic symmetry capability. [application to solid rocket propellant grains and space antennas

    NASA Technical Reports Server (NTRS)

    Macneal, R. H.; Harder, R. L.; Mason, J. B.

    1973-01-01

    A development for NASTRAN which facilitates the analysis of structures made up of identical segments symmetrically arranged with respect to an axis is described. The key operation in the method is the transformation of the degrees of freedom for the structure into uncoupled symmetrical components, thereby greatly reducing the number of equations which are solved simultaneously. A further reduction occurs if each segment has a plane of reflective symmetry. The only required assumption is that the problem be linear. The capability, as developed, will be available in level 16 of NASTRAN for static stress analysis, steady state heat transfer analysis, and vibration analysis. The paper includes a discussion of the theory, a brief description of the data supplied by the user, and the results obtained for two example problems. The first problem concerns the acoustic modes of a long prismatic cavity imbedded in the propellant grain of a solid rocket motor. The second problem involves the deformations of a large space antenna. The latter example is the first application of the NASTRAN Cyclic Symmetry capability to a really large problem.

  19. 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.

  20. Feasibility of an advanced thrust termination assembly for a solid propellant rocket motor

    NASA Technical Reports Server (NTRS)

    1975-01-01

    A total of 68 quench tests were conducted in a vented bomb assembly (VBA). Designed to simulate full-scale motor operating conditions, this laboratory apparatus uses a 2-inch-diameter, end-burning propellant charge and an insulated disc of consolidated hydrated aluminum sulfate along with the explosive charge necessary to disperse the salt and inject it onto the burning surface. The VBA was constructed to permit variation of motor design parameters of interest; i.e., weight of salt per unit burning surface area, weight of explosive per unit weight of salt, distance from salt surface to burning surface, incidence angle of salt injection, chamber pressure, and burn time. Completely satisfactory salt quenching, without re-ignition, occurred in only two VBA tests. These were accomplished with a quench charge ratio (QCR) of 0.023 lb salt per square inch of burning surface at dispersing charge ratios (DCR) of 13 and 28 lb of salt per lb of explosive. Candidate materials for insulating salt charges from the rocket combustion environment were evaluated in firings of 5-inch-diameter, uncured end-burner motors. A pressed, alumina ceramic fiber material was selected for further evaluation and use in the final demonstration motor.

  1. The measurement, modeling, and prediction of traction for rocket propellant 1

    NASA Technical Reports Server (NTRS)

    Tevaarwerk, J. L.

    1989-01-01

    Traction tests were performed on RP-1, a common kerosene based rocket propellant. Traction data on this fluid are required for purposes of turbopump bearing design, using codes such as SHABERTH. To obtain the traction data, an existing twin disc machine was used, operating under the side slip mode and using elliptical contacts. The range of test variables were: contact peak Hertz stress from 1.0 to 2.0 GPa, disc surface speed from 10 to 50 m/s, fluid inlet temperature from 30 to 70 C, and with a contact aspect ratio of 1.7. The resulting traction curves were reduced to fundamental fluid property parameters using the Johnson and Tevaarwerk traction model. Theoretical traction predictions were performed by back substitution of the fundamental properties into the traction model. Comparison of the predicted with the measured curves gives a high degree of confidence in the correctness of the traction model. For purposes of input to the NASA SHABERTH program, the traction model was next used to predict the expected traction of RP-1 under line contact conditions.

  2. Baseline Computational Fluid Dynamics Methodology for Longitudinal-Mode Liquid-Propellant Rocket Combustion Instability

    NASA Technical Reports Server (NTRS)

    Litchford, R. J.

    2005-01-01

    A computational method for the analysis of longitudinal-mode liquid rocket combustion instability has been developed based on the unsteady, quasi-one-dimensional Euler equations where the combustion process source terms were introduced through the incorporation of a two-zone, linearized representation: (1) A two-parameter collapsed combustion zone at the injector face, and (2) a two-parameter distributed combustion zone based on a Lagrangian treatment of the propellant spray. The unsteady Euler equations in inhomogeneous form retain full hyperbolicity and are integrated implicitly in time using second-order, high-resolution, characteristic-based, flux-differencing spatial discretization with Roe-averaging of the Jacobian matrix. This method was initially validated against an analytical solution for nonreacting, isentropic duct acoustics with specified admittances at the inflow and outflow boundaries. For small amplitude perturbations, numerical predictions for the amplification coefficient and oscillation period were found to compare favorably with predictions from linearized small-disturbance theory as long as the grid exceeded a critical density (100 nodes/wavelength). The numerical methodology was then exercised on a generic combustor configuration using both collapsed and distributed combustion zone models with a short nozzle admittance approximation for the outflow boundary. In these cases, the response parameters were varied to determine stability limits defining resonant coupling onset.

  3. Experimental study of a valveless pulse detonation rocket engine using nontoxic hypergolic propellants

    NASA Astrophysics Data System (ADS)

    Kan, Brandon K.

    A pulsed detonation rocket engine concept was explored through the use of hypergolic propellants in a fuel-centered pintle injector combustor. The combustor design yielded a simple open ended chamber with a pintle type injection element and pressure instrumentation. High-frequency pressure measurements from the first test series showed the presence of large pressure oscillations in excess of 2000 psia at frequencies between 400-600 hz during operation. High-speed video confirmed the high-frequency pulsed behavior and large amounts of after burning. Damaged hardware and instrumentation failure limited the amount of data gathered in the first test series, but the experiments met original test objectives of producing large over-pressures in an open chamber. A second test series proceeded by replacing hardware and instrumentation, and new data showed that pulsed events produced under expanded exhaust prior to pulsing, peak pressures around 8000 psi, and operating frequencies between 400-800 hz. Later hot-fires produced no pulsed behavior despite undamaged hardware. The research succeeded in producing pulsed combustion behavior using hypergolic fuels in a pintle injector setup and provided insights into design concepts that would assist future injector designs and experimental test setups.

  4. Cooling Duct Analysis for Transpiration/Film Cooled Liquid Propellant Rocket Engines

    NASA Technical Reports Server (NTRS)

    Micklow, Gerald J.

    1996-01-01

    The development of a low cost space transportation system requires that the propulsion system be reusable, have long life, with good performance and use low cost propellants. Improved performance can be achieved by operating the engine at higher pressure and temperature levels than previous designs. Increasing the chamber pressure and temperature, however, will increase wall heating rates. This necessitates the need for active cooling methods such as film cooling or transpiration cooling. But active cooling can reduce the net thrust of the engine and add considerably to the design complexity. Recently, a metal drawing process has been patented where it is possible to fabricate plates with very small holes with high uniformity with a closely specified porosity. Such a metal plate could be used for an inexpensive transpiration/film cooled liner to meet the demands of advanced reusable rocket engines, if coolant mass flow rates could be controlled to satisfy wall cooling requirements and performance. The present study investigates the possibility of controlling the coolant mass flow rate through the porous material by simple non-active fluid dynamic means. The coolant will be supplied to the porous material by series of constant geometry slots machined on the exterior of the engine.

  5. 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.

  6. Design and Fabrication of a 200N Thrust Rocket Motor Based on NH4ClO4+Al+HTPB as Solid Propellant

    NASA Astrophysics Data System (ADS)

    Wahid, Mastura Ab; Ali, Wan Khairuddin Wan

    2010-06-01

    The development of rocket motor using potassium nitrate, carbon and sulphur mixture has successfully been developed by researchers and students from UTM and recently a new combination for solid propellant is being created. The new solid propellant will combine a composition of Ammonium perchlorate, NH4ClO4 with aluminium, Al and Hydroxyl Terminated Polybutadiene, HTPB as the binder. It is the aim of this research to design and fabricate a new rocket motor that will produce a thrust of 200N by using this new solid propellant. A static test is done to obtain the thrust produced by the rocket motor and analyses by observation and also calculation will be done. The experiment for the rocket motor is successful but the thrust did not achieve its required thrust.

  7. Cooling of in-situ propellant rocket engines for Mars mission. M.S. Thesis - Cleveland State Univ.

    NASA Technical Reports Server (NTRS)

    Armstrong, Elizabeth S.

    1991-01-01

    One propulsion option of a Mars ascent/descent vehicle is multiple high-pressure, pump-fed rocket engines using in-situ propellants, which have been derived from substances available on the Martian surface. The chosen in-situ propellant combination for this analysis is carbon monoxide as the fuel and oxygen as the oxidizer. Both could be extracted from carbon dioxide, which makes up 96 percent of the Martian atmosphere. A pump-fed rocket engine allows for higher chamber pressure than a pressure-fed engine, which in turn results in higher thrust and in higher heat flux in the combustion chamber. The heat flowing through the wall cannot be sufficiently dissipated by radiation cooling and, therefore, a regenerative coolant may be necessary to avoid melting the rocket engine. The two possible fluids for this coolant scheme, carbon monoxide and oxygen, are compared analytically. To determine their heat transfer capability, they are evaluated based upon their heat transfer and fluid flow characteristics.

  8. High performance N2O4/amine elements. [propellant tests of hypergolic rocket propellants used in Space Shuttle Orbiters

    NASA Technical Reports Server (NTRS)

    Falk, A. Y.

    1976-01-01

    An analytical and experimental investigation was conducted to develop an understanding of the mechanisms that cause reactive stream separation, commonly called blowapart, for hypergolic propellants. The investigation was limited to a N2O4/MMH propellant combination and to a range of engine-operating conditions applicable to the space tug and space shuttle attitude control and orbital maneuvering engines. Primary test variables were: chamber pressure (1 to 20 atm), fuel injection temperature (283 to 400 K)m and propellant injection velocity (9 to 50 m/s). The injector configuration studied was the unlike doublet. The reactive stream separation experiments were conducted using special combustors designed to permit photography of the near-injector spray combustion flow field. Analysis of color motion pictures provided the means of determining the occurrence of reactive stream separation.

  9. 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.

  10. Space shuttle SRM plume expansion sensitivity analysis. [flow characteristics of exhaust gases from solid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Smith, S. D.; Tevepaugh, J. A.; Penny, M. M.

    1975-01-01

    The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.

  11. Analytical Investigation of the Effect of Turbopump Design on Gross-Weight Characteristics of a Hydrogen-Propelled Nuclear Rocket

    NASA Technical Reports Server (NTRS)

    Rohlik, Harold E.; Crouse, James E.

    1959-01-01

    The effect of turbopump design on rocket gross weight was investigated for a high-pressure bleed-type hydrogen-reactor long-range rocket with a fixed mission. Axial-flow, mixed-flow, and centrifugal pumps driven by single and twin turbines were considered. With an efficiency of 0.7 assumed for all pumps, the lowest rocket gross weights were obtained with an axial-flow or a mixed-flow pump driven by a single turbine of at least eight stages. All turbopump combinations could be used, however, with gross weight varying less than 8 percent for a given payload. Turbopump efficiencies have a significant effect on the ratio of gross weight to payload with the magnitude of the effect determined by the ratio of rocket structural weight to total propellant weight. One point in pump efficiency is worth 0.2 percent in gross weight for a given payload with a structural weight parameter of 0.1 and 0.6 percent with a structural weight parameter of 0.2. Turbine and pump weights are much less significant in terms of gross-to-pay weight ratio than the efficiencies of these components. One point in pump efficiency is equivalent to approximately 13 percent in pump weight, while 1 point in turbine efficiency is equivalent to about 7 percent in turbine weight.

  12. Concerning the problem of dynamic damping of the vibration combustion self-oscillations in a liquid-propellant rocket engine

    NASA Astrophysics Data System (ADS)

    Basok, B. I.; Gotsulenko, V. V.; Gotsulenko, V. N.

    2012-11-01

    The reason for the decrease in the amplitude of longitudinal vibration combustion self-oscillations in the combustion chamber of a liquid-propellant rocket engine by means of antipulse partitions has been justified. A mathematical model of the development of combustion instability in such a chamber on attachment of a Helmholtz resonator to it has been obtained. The character of the damping of vibration combustion self-oscillations excited by the action of the Crocco mechanisms and negative thermal resistance, when varying the acoustic parameters of the resonator and of the pressure head characteristics of combustion chamber is established.

  13. Metallized solid rocket propellants based on AN/AP and PSAN/AP for access to space

    NASA Astrophysics Data System (ADS)

    Levi, S.; Signoriello, D.; Gabardi, A.; Molinari, M.; Galfetti, L.; Deluca, L. T.; Cianfanelli, S.; Klyakin, G. F.

    2009-09-01

    Solid rocket propellants based on dual mixes of inorganic crystalline oxidizers (ammonium nitrate (AN) and ammonium perchlorate (AP)) with binder and a mixture of micrometric-nanometric aluminum were investigated. Ammonium nitrate is a low-cost oxidizer, producing environment friendly combustion products but with lower specific impulse compared to AP. The better performance obtained with AP and the low quantity of toxic emissions obtained by using AN have suggested an interesting compromise based on a dual mixture of the two oxidizers. To improve the thermal response of raw AN, different types of phase stabilized AN (PSAN) and AN/AP co-crystals were investigated.

  14. Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft

    NASA Technical Reports Server (NTRS)

    Werka, Robert; Clark, Rod; Sheldon, Rob; Percy, Tom

    2012-01-01

    The March, 2012 issue of Aerospace America stated that ?the near-to-medium prospects for applying advanced propulsion to create a new era of space exploration are not very good. In the current world, we operate to the Moon by climbing aboard a Carnival Cruise Lines vessel (Saturn 5), sail from the harbor (liftoff) shedding whole decks of the ship (staging) along the way and, having reached the return leg of the journey, sink the ship (burnout) and return home in a lifeboat (Apollo capsule). Clearly this is an illogical way to travel, but forced on Explorers by today's propulsion technology. However, the article neglected to consider the one propulsion technology, using today's physical principles that offer continuous, substantial thrust at a theoretical specific impulse of 1,000,000 sec. This engine unequivocally can create a new era of space exploration that changes the way spacecraft operate. Today's space Explorers could travel in Cruise Liner fashion using the technology not considered by Aerospace America, the novel Dusty Plasma Fission Fragment Rocket Engine (FFRE). This NIAC study addresses the FFRE as well as its impact on Exploration Spacecraft design and operation. It uses common physics of the relativistic speed of fission fragments to produce thrust. It radiatively cools the fissioning dusty core and magnetically controls the fragments direction to practically implement previously patented, but unworkable designs. The spacecraft hosting this engine is no more complex nor more massive than the International Space Station (ISS) and would employ the successful ISS technology for assembly and check-out. The elements can be lifted in "chunks" by a Heavy Lift Launcher. This Exploration Spacecraft would require the resupply of small amounts of nuclear fuel for each journey and would be an in-space asset for decades just as any Cruise Liner on Earth. This study has synthesized versions of the FFRE, integrated one concept onto a host spacecraft designed for

  15. Aluminum/hydrocarbon gel propellants: An experimental and theoretical investigation of secondary atomization and predicted rocket engine performance

    NASA Astrophysics Data System (ADS)

    Mueller, Donn Christopher

    1997-12-01

    Experimental and theoretical investigations of aluminum/hydrocarbon gel propellant secondary atomization and its potential effects on rocket engine performance were conducted. In the experimental efforts, a dilute, polydisperse, gel droplet spray was injected into the postflame region of a burner and droplet size distributions was measured as a function of position above the burner using a laser-based sizing/velocimetry technique. The sizing/velocimetry technique was developed to measure droplets in the 10-125 mum size range and avoids size-biased detection through the use of a uniformly illuminated probe volume. The technique was used to determine particle size distributions and velocities at various axial locations above the burner for JP-10, and 50 and 60 wt% aluminum gels. Droplet shell formation models were applied to aluminum/hydrocarbon gels to examine particle size and mass loading effects on the minimum droplet diameter that will permit secondary atomization. This diameter was predicted to be 38.1 and 34.7 mum for the 50 and 60 wt% gels, which is somewhat greater than the experimentally measured 30 and 25 mum diameters. In the theoretical efforts, three models were developed and an existing rocket code was exercised to gain insights into secondary atomization. The first model was designed to predict gel droplet properties and shell stresses after rigid shell formation, while the second, a one-dimensional gel spray combustion model was created to quantify the secondary atomization process. Experimental and numerical comparisons verify that secondary atomization occurs in 10-125 mum diameter particles although an exact model could not be derived. The third model, a one-dimensional gel-fueled rocket combustion chamber, was developed to evaluate secondary atomization effects on various engine performance parameters. Results show that only modest secondary atomization may be required to reduce propellant burnout distance and radiation losses. A solid propellant

  16. The starting transient of solid propellant rocket motors with high internal gas velocities. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Peretz, A.; Caveny, L. H.; Kuo, K. K.; Summerfield, M.

    1973-01-01

    A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

  17. Educating Tomorrow's Aerrospace Engineers by Developing and Launching Liquid-Propelled Rockets

    NASA Astrophysics Data System (ADS)

    Besnard, Eric; Garvey, John; Holleman, Tom; Mueller, Tom

    2002-01-01

    conducted at California State University, Long Beach (CSULB), in which engineering students develop and launch liquid propelled rockets. The program is articulated around two main activities, each with specific objectives. The first component of CALVEIN is a systems integration laboratory where students develop/improve vehicle subsystems and integrate them into a vehicle (Prospector-2 - P-2 - for the 2001-02 academic year - AY). This component has three main objectives: (1) Develop hands- on skills for incoming students and expose them to aerospace hardware; (2) allow for upper division students who have been involved in the program to mentor incoming students and manage small teams; and (3) provide students from various disciplines within the College of Engineering - and other universities - with the chance to develop/improve subsystems on the vehicle. Among recent student projects conducted as part of this component are: a new 1000 lbf thrust engine using pintle injector technology, which was successfully tested on Dec. 1, 2001 and flown on Prospector-2 in Feb. 2002 (developed by CSULB Mechanical and Aerospace Engineering students); a digital flight telemetry package (developed by CSULB Electrical Engineering students); a new recovery system where a mechanical system replaces pyrotechnics for parachute release (developed by CSULB Mechanical and Aerospace Engineering students); and a 1-ft payload bay to accommodate experimental payloads (e.g. "CANSATS" developed by Stanford University students). The second component of CALVEIN is a formal Aerospace System Design curriculum. In the first-semester, from top-level system requirements, the students perform functional analysis, define the various subsystems and derive their requirements. These are presented at the Systems Functional and Requirement Reviews (SFR &SRR). The methods used for validation and verification are determined. Specifications and Interface Control Documents (ICD) are generated by the student team

  18. Propellant grain dynamics in aft attach ring of shuttle solid rocket booster

    NASA Technical Reports Server (NTRS)

    Verderaime, V.

    1979-01-01

    An analytical technique for implementing simultaneously the temperature, dynamic strain, real modulus, and frequency properties of solid propellant in an unsymmetrical vibrating ring mode is presented. All dynamic parameters and sources are defined for a free vibrating ring-grain structure with initial displacement and related to a forced vibrating system to determine the change in real modulus. Propellant test data application is discussed. The technique was developed to determine the aft attach ring stiffness of the shuttle booster at lift-off.

  19. Characterization of the exhaust particulates in the ground cloud and high-altitude plume of large solid-propellant booster rockets

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Bowyer, J. M.; Varsi, G.; Laue, E. G.; Gauldin, R.

    1980-01-01

    The report is concerned with the characterization of Al2O3 particles in the atmosphere. These particles comprise one of the major combustion products of the rocket propellant employed in the Space Shuttle boosters. A ground cloud and stratospheric plume are considered. It is concluded that the typical residence times in the atmosphere are much longer than earlier estimates have indicated.

  20. Numerical investigation of the influence of crystallization of ultrafine particles of aluminum oxide on energy characteristics of solid-propellant rocket engine

    NASA Astrophysics Data System (ADS)

    Dyachenko, N. N.; Dyachenko, L. I.

    2014-08-01

    The results of numerical investigation of a multiphase flow considering coagulation, crushing and crystallization of the particles of polydispersed condensate in the nozzles of solid-propellant rocket engine are presented. The influence of particles crystallization on the energy characteristics of the engine is shown.

  1. Removing hydrochloric acid exhaust products from high performance solid rocket propellant using aluminum-lithium alloy.

    PubMed

    Terry, Brandon C; Sippel, Travis R; Pfeil, Mark A; Gunduz, I Emre; Son, Steven F

    2016-11-01

    Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solid propellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (ISP). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal ISP by ∼7s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solid propellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5±4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption. PMID:27289269

  2. Analysis of Flow-System Starting Dynamics of Turbopump-Fed Liquid-Propellant Rocket

    NASA Technical Reports Server (NTRS)

    Krebs, Richard P.; Hart, Clint E.

    1959-01-01

    Two rocket configurations with turbopump drive were investigated analytically. In one configuration the inlet pressure to the turbine was fixed at the design value. The second configuration employed a "bootstrap" technique for supplying energy to the turbine. An injector was the chief resistance between the pump and the rocket combustion chamber. From the analysis two parameters were developed from which the speed response time of the turbopump, the flow response time, and the maximum dynamic line loss could be evaluated. These parameters were functions of turbopump moment of inertia, design performance of the turbine, and flow-system geometry. The moment of inertia of the turbopump and the ratio of turbine torque at zero speed to design torque had the most influence on the starting dynamics of the flow system. These parameters were also applicable to the bootstrap configuration as long as the inlet pressure to the turbine exceeded half the design value.

  3. Uncertainties in the characterization of the thermal environment of a solid rocket propellant fire

    SciTech Connect

    Diaz, J.C.

    1993-10-01

    There has been an interest in developing models capable of predicting the response of systems to Minuteman (MM) III third-stage solid propellant fires. Input parameters for such an effort include the boundary conditions that describe the fire temperature, heat flux, emissivity, and propellant burn rate. In this study scanning spectroscopy and pyrometry were used to infer plume temperatures. Each diagnostic system possessed strengths and weaknesses. The intention was to use various supportive methods to infer plume temperature and emissivity, because no one diagnostic had proven capabilities for determining temperature under these conditions. Furthermore, these diagnostics were being used near the limit of their applicability. All these points created some uncertainty in the data collected.

  4. Distributed Parameter Analysis of Pressure and Flow Disturbances in Rocket Propellant Feed Systems

    NASA Technical Reports Server (NTRS)

    Dorsch, Robert G.; Wood, Don J.; Lightner, Charlene

    1966-01-01

    A digital distributed parameter model for computing the dynamic response of propellant feed systems is formulated. The analytical approach used is an application of the wave-plan method of analyzing unsteady flow. Nonlinear effects are included. The model takes into account locally high compliances at the pump inlet and at the injector dome region. Examples of the calculated transient and steady-state periodic responses of a simple hypothetical propellant feed system to several types of disturbances are presented. Included are flow disturbances originating from longitudinal structural motion, gimbaling, throttling, and combustion-chamber coupling. The analytical method can be employed for analyzing developmental hardware and offers a flexible tool for the calculation of unsteady flow in these systems.

  5. PC programs for the prediction of the linear stability behavior of liquid propellant propulsion systems and application to current MSFC rocket engine test programs, volume 1

    NASA Technical Reports Server (NTRS)

    Doane, George B., III; Armstrong, W. C.

    1990-01-01

    Research on propulsion stability (chugging and acoustic modes), and propellant valve control was investigated. As part of the activation of the new liquid propulsion test facilities, it is necessary to analyze total propulsion system stability. To accomplish this, several codes were built to run on desktop 386 machines. These codes enable one to analyze the stability question associated with the propellant feed systems. In addition, further work was adapted to this computing environment and furnished along with other codes. This latter inclusion furnishes those interested in high frequency oscillatory combustion behavior (that does not couple to the feed system) a set of codes for study of proposed liquid rocket engines.

  6. A General Method for Automatic Computation of Equilibrium Compositions and Theoretical Rocket Performance of Propellants

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Zeleznik, Frank J.; Huff, Vearl N.

    1959-01-01

    A general computer program for chemical equilibrium and rocket performance calculations was written for the IBM 650 computer with 2000 words of drum storage, 60 words of high-speed core storage, indexing registers, and floating point attachments. The program is capable of carrying out combustion and isentropic expansion calculations on a chemical system that may include as many as 10 different chemical elements, 30 reaction products, and 25 pressure ratios. In addition to the equilibrium composition, temperature, and pressure, the program calculates specific impulse, specific impulse in vacuum, characteristic velocity, thrust coefficient, area ratio, molecular weight, Mach number, specific heat, isentropic exponent, enthalpy, entropy, and several thermodynamic first derivatives.

  7. A Rocket Engine for Mars Sample Return Using In Situ Propellants

    NASA Technical Reports Server (NTRS)

    Linne, Diane L.

    1997-01-01

    Recent studies for the planned Mars sample return mission were reviewed and modified to utilize carbon monoxide and oxygen as potential in situ propellants. Based on these studies a representative full scale engine thrust of 2225 N (500 lbf) was selected as appropriate to demonstrate performance, and the design for that engine is presented. Previous experimental results combined with parametric analyses were used to define the geometry for the engine which operates on liquid carbon monoxide and liquid oxygen. The engine was constructed using a combination of high-temperature alloys and lightweight ceramics. The materials selected were hafnium oxide, iridium, rhenium, and carbon-carbon.

  8. Liquid propellant rocket engine combustion simulation with a time-accurate CFD method

    NASA Technical Reports Server (NTRS)

    Chen, Y. S.; Shang, H. M.; Liaw, Paul; Hutt, J.

    1993-01-01

    Time-accurate computational fluid dynamics (CFD) algorithms are among the basic requirements as an engineering or research tool for realistic simulations of transient combustion phenomena, such as combustion instability, transient start-up, etc., inside the rocket engine combustion chamber. A time-accurate pressure based method is employed in the FDNS code for combustion model development. This is in connection with other program development activities such as spray combustion model development and efficient finite-rate chemistry solution method implementation. In the present study, a second-order time-accurate time-marching scheme is employed. For better spatial resolutions near discontinuities (e.g., shocks, contact discontinuities), a 3rd-order accurate TVD scheme for modeling the convection terms is implemented in the FDNS code. Necessary modification to the predictor/multi-corrector solution algorithm in order to maintain time-accurate wave propagation is also investigated. Benchmark 1-D and multidimensional test cases, which include the classical shock tube wave propagation problems, resonant pipe test case, unsteady flow development of a blast tube test case, and H2/O2 rocket engine chamber combustion start-up transient simulation, etc., are investigated to validate and demonstrate the accuracy and robustness of the present numerical scheme and solution algorithm.

  9. Performance of a UTC FW-4S solid propellant rocket motor under the command effects of simulated altitude and rotational spin

    NASA Technical Reports Server (NTRS)

    Merryman, H. L.; Smith, L. R.

    1974-01-01

    One United Technology Center FW-4S solid-propellant rocket motor was fired at an average simulated altitude of 103,000 ft while spinning about its axial centerline at 180 rpm. The objectives of the test program were to determine motor altitude ballistic performance including the measurement of the nonaxial thrust vector and to demonstrate structural integrity of the motor case and nozzle. These objectives are presented and discussed.

  10. Flow visualization of a rocket injector spray using gelled propellant simulants

    NASA Technical Reports Server (NTRS)

    Green, James M.; Rapp, Douglas C.; Roncace, James

    1991-01-01

    A study was conducted at NASA-Lewis to compare the atomization characteristics of gelled and nongelled propellant simulants. A gelled propellant simulant composed of water, sodium hydroxide, and an acrylic acid polymer resin (as the gelling agent) was used to simulate the viscosity of an aluminum/PR-1 metallized fuel gel. Water was used as a comparison fluid to isolate the rheological effects of the water-gel and to simulate nongelled RP-1. The water-gel was injected through the central orifice of a triplet injector element and the central post of a coaxial injector element. Nitrogen gas flowed through the outer orifices of the triplet injector element and through the annulus of the coaxial injector element and atomized the gelled and nongelled liquids. Photographs of the water-gel spray patterns at different operating conditions were compared with images obtained using water and nitrogen. A laser light was used for illumination of the sprays. The results of the testing showed that the water sprays produced a finer and more uniform atomization than the water-gel sprays. Rheological analysis of the water-gel showed poor atomization caused by high viscosity of water-gel delaying the transition to turbulence.

  11. Measurement of Solid Rocket Propellant Burning Rate Using X-ray Imaging

    NASA Astrophysics Data System (ADS)

    Denny, Matthew D.

    The burning rate of solid propellants can be difficult to measure for unusual burning surface geometries, but X-ray imaging can be used to measure burning rate. The objectives of this work were to measure the baseline burning rate of an electrically-controlled solid propellant (ESP) formulation with real-time X-ray radiography and to determine the uncertainty of the measurements. Two edge detection algorithms were written to track the burning surface in X-ray videos. The edge detection algorithms were informed by intensity profiles of simulated 2-D X-ray images. With a 95% confidence level, the burning rates measured by the Projected-Slope Intersection algorithm in the two combustion experiments conducted were 0.0839 in/s +/-2.86% at an average pressure of 407 psi +/-3.6% and 0.0882 in/s +/-3.04% at 410 psi +/-3.9%. The uncertainty percentages were based on the statistics of a Monte Carlo analysis on burning rate.

  12. Development of the platelet micro-orifice injector. [for liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    La Botz, R. J.

    1984-01-01

    For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.

  13. Probabilistic low cycle fatigue failure analysis with application to liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Newlin, L.; Sutharshana, S.; Ebbeler, D.; Moore, N.; Fox, E.

    1990-01-01

    A probabilistic Low Cycle Fatigue (LCF) failure analysis of a candidate turbine disk for use in a turbopump of a rocket engine of the Space Shuttle Main Engine class is described. A state-of-the-art LCF failure prediction method was used in a Monte Carlo simulation to generate a distribution of failure lives. A stochastic Stress/Life (S/N) model was used for materials characterization. The LCF failure model expresses fatigue life as a function of stochastic parameters including environmental parameters, loads, material properties, structural parameters, and model specification errors. The rationale for the particular characterization of each stochastic input parameter is described. The results and interpretation of the failure analysis are given.

  14. Probabilistic high cycle fatigue failure analysis with application to liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Sutharshana, S.; Newlin, L.; Ebbeler, D.; Moore, N.; O'Hara, K.

    1990-01-01

    A probabilistic high cycle fatigue (HCF) failure analysis of a welded duct in a rocket engine of the Space Shuttle main engine class is described. A state-of-the-art HCF failure prediction method was used in a Monte Carlo simulation to generate a distribution of failure lives. A stochastic stress/life model is used for material characterization, and a composite stress history is generated for accurately deriving the stress cycles for the fatigue-damage calculations. The HCF failure model expresses fatigue life as a function of stochastic parameters including environment, loads, material properties, geometry, and model specification errors. A series of HCF failure life analyses were performed to study the impact of a fixed parameter and to assess the importance of each stochastic input parameter through marginal analyses.

  15. Conversion of the rocket propellant UDMH to a reagent useful in vicarious nucleophilic substitution reactions

    SciTech Connect

    Mitchell, A.R.; Pagoria, P.F.; Schmidt, R.D.

    1995-11-10

    The objective of our program is to develop novel, innovative solutions for the disposal of surplus energetic materials resulting from the demilitarization of conventional and nuclear munitions. In this report we describe the use of surplus propellant (UDMH) and explosives (TNT, Explosive D) as chemical precursors for higher value products. The conversion of UDMH to 1,1,1-trimethylhydrazinium iodide (TMHI) provides a new aminating reagent for use in Vicarious Nucleophilic Substitution (VNS) reactions. When TMHI is reacted with various nitroarenes the amino functionality is introduced in good to excellent yields. Thus, 2,4,6-trinitroaniline (picramide) reacts with TMHI to give 1,3,5-triamino-2,4,6-trinitroaniline (TATB) while 2,4,6-trinitrotoluene (TNT) reacts with TMHI to give 3,5-diamino-2,4,6-trinitrotoluene (DATNT). The advantages, scope and limitations of the VNS approach and the use of TMHI are discussed.

  16. Nuclear Magnetic Resonance Imaging of Solid Rocket Propellants at 14.1 T

    NASA Astrophysics Data System (ADS)

    Maas, W. E.; Merwin, L. H.; Cory, D. G.

    1997-11-01

    Proton NMR images of solid propellant materials, consisting of a polybutadiene binder material filled with 82% solid particles, have been obtained at a magnetic field strength of 14.1 T and at a resolution of 8.5 × 8.5 μm. The images are the first of elastomeric materials obtained at a proton frequency of 600 MHz and have the highest spatial resolution yet reported. The images display a high contrast and are rich in information content. They reveal the distribution of individual filler particles in the polymer matrix as well as a thin polymer film of about 10-30 μm which is found to surround some of the larger filler particles.

  17. Uncertainty Quantification of Non-linear Oscillation Triggering in a Multi-injector Liquid-propellant Rocket Combustion Chamber

    NASA Astrophysics Data System (ADS)

    Popov, Pavel; Sideris, Athanasios; Sirignano, William

    2014-11-01

    We examine the non-linear dynamics of the transverse modes of combustion-driven acoustic instability in a liquid-propellant rocket engine. Triggering can occur, whereby small perturbations from mean conditions decay, while larger disturbances grow to a limit-cycle of amplitude that may compare to the mean pressure. For a deterministic perturbation, the system is also deterministic, computed by coupled finite-volume solvers at low computational cost for a single realization. The randomness of the triggering disturbance is captured by treating the injector flow rates, local pressure disturbances, and sudden acceleration of the entire combustion chamber as random variables. The combustor chamber with its many sub-fields resulting from many injector ports may be viewed as a multi-scale complex system wherein the developing acoustic oscillation is the emergent structure. Numerical simulation of the resulting stochastic PDE system is performed using the polynomial chaos expansion method. The overall probability of unstable growth is assessed in different regions of the parameter space. We address, in particular, the seven-injector, rectangular Purdue University experimental combustion chamber. In addition to the novel geometry, new features include disturbances caused by engine acceleration and unsteady thruster nozzle flow.

  18. Solid-propellant rocket motor internal ballistic performance variation analysis, phase 2

    NASA Technical Reports Server (NTRS)

    Sforzini, R. H.; Foster, W. A., Jr.

    1976-01-01

    The Monte Carlo method was used to investigate thrust imbalance and its first time derivative throughtout the burning time of pairs of solid rocket motors firing in parallel. Results obtained compare favorably with Titan 3 C flight performance data. Statistical correlations of the thrust imbalance at various times with corresponding nominal trace slopes suggest several alternative methods of predicting thrust imbalance. The effect of circular-perforated grain deformation on internal ballistics is discussed, and a modified design analysis computer program which permits such an evaluation is presented. Comparisons with SRM firings indicate that grain deformation may account for a portion of the so-called scale factor on burning rate between large motors and strand burners or small ballistic test motors. Thermoelastic effects on burning rate are also investigated. Burning surface temperature is calculated by coupling the solid phase energy equation containing a strain rate term with a model of gas phase combustion zone using the Zeldovich-Novozhilov technique. Comparisons of solutions with and without the strain rate term indicate a small but possibly significant effect of the thermoelastic coupling.

  19. Effect of Nozzle Nonlinearities upon Nonlinear Stability of Liquid Propellant Rocket Motors

    NASA Technical Reports Server (NTRS)

    Padmanabhan, M. S.; Powell, E. A.; Zinn, B. T.

    1975-01-01

    A three dimensional, nonlinear nozzle admittance relation is developed by solving the wave equation describing finite amplitude oscillatory flow inside the subsonic portion of a choked, slowly convergent axisymmetric nozzle. This nonlinear nozzle admittance relation is then used as a boundary condition in the analysis of nonlinear combustion instability in a cylindrical liquid rocket combustor. In both nozzle and chamber analyses solutions are obtained using the Galerkin method with a series expansion consisting of the first tangential, second tangential, and first radial modes. Using Crocco's time lag model to describe the distributed unsteady combustion process, combustion instability calculations are presented for different values of the following parameters: (1) time lag, (2) interaction index, (3) steady-state Mach number at the nozzle entrance, and (4) chamber length-to-diameter ratio. In each case, limit cycle pressure amplitudes and waveforms are shown for both linear and nonlinear nozzle admittance conditions. These results show that when the amplitudes of the second tangential and first radial modes are considerably smaller than the amplitude of the first tangential mode the inclusion of nozzle nonlinearities has no significant effect on the limiting amplitude and pressure waveforms.

  20. Asbestos Free Insulation Development for the Space Shuttle Solid Propellant Rocket Motor (RSRM)

    NASA Technical Reports Server (NTRS)

    Allred, Larry D.; Eddy, Norman F.; McCool, A. A. (Technical Monitor)

    2000-01-01

    Asbestos has been used for many years as an ablation inhibitor in insulating materials. It has been a constituent of the AS/NBR insulation used to protect the steel case of the RSRM (Reusable Solid Rocket Motor) since its inception. This paper discusses the development of a potential replacement RSRM insulation design, several of the numerous design issues that were worked and processing problems that were resolved. The earlier design demonstration on FSM-5 (Flight Support Motor) of the selected 7% and 11% Kevlar(registered) filled EPDM (KF/EPDM) candidate materials was expanded. Full-scale process simulation articles were built and FSM-8 was manufactured using multiple Asbestos Free (AF) components and materials. Two major problems had to be overcome in developing the AF design. First, bondline corrosion, which occurred in the double-cured region of the aft dome, had to be eliminated. Second, KF/EPDM creates high levels of electrostatic energy (ESE), which does not readily dissipate from the insulation surface. An uncontrolled electrostatic discharge (ESD) of this surface energy during many phases of production could create serious safety hazards. Numerous processing changes were implemented and a conductive paint was developed to prevent exposed external insulation surfaces from generating ESE/ESD. Additionally, special internal instrumentation was incorporated into FSM-8 to record real-time internal motor environment data. These data included inhibitor insulation erosion rates and internal thermal environments. The FSM-8 static test was successfully conducted in February 2000 and much valuable data were obtained to characterize the AF insulation design.

  1. The dynamics of a gas-dust cloud expansion in the upper atmosphere at a shutdown of solid-propellant rocket engines

    NASA Astrophysics Data System (ADS)

    Nikolaishvili, S. Sh.; Platov, Yu. V.; Chernouss, S. A.

    2015-09-01

    The velocity of spherical gas-dust cloud expansion in the situation when the stages of solid-propellant rocket separate in the upper atmosphere have been determined. The measured velocity vary from 2.5 to 7.5 km/s. The dispersed component accelerates at the front of a shock that develops at engine-thrust shutdown. The model calculations of the gas-dust cloud luminosity intensity qualitatively coincide with the photometric profiles of object images. Such formations can vary from almost homogeneous ball-shaped clouds to rather thin spherical shells depending on the gas-dust cloud mass and the matter distribution within this cloud.

  2. 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.

  3. Effect of Propellant and Catalyst Bed Temperatures on Thrust Buildup in Several Hydrogen Peroxide Reaction Control Rockets

    NASA Technical Reports Server (NTRS)

    Wanhainen, John P.; Ross, Phil S.; DeWitt, Richard L.

    1960-01-01

    An investigation was undertaken to determine the effect of chamber and propellant feed temperatures on the starting characteristics of hydrogen peroxide thrust chambers. Start delay times for two types of thrust chamber designs in the 1- to 24-pound-thrust range were obtained over a range of chamber and propellant feed temperatures from 30 to 100 F. Start delay times obtained during the first minute of catalyst bed life and again after 6 minutes of total accumulated running time are presented as a function of chamber and propellant feed temperatures. The initial cold-start delay time of the hydrogen peroxide thrust chambers investigated was approximately 0.150 second to attain 90 percent of steady-state chamber pressure at chamber and propellant feed temperatures of 70 F and above. Both thrust chamber designs could be started at chamber and propellant feed temperatures as low as 30 F; start delay times did, however, generally increase at low temperatures. When the chamber was at an elevated temperature from a preceding firing, the start delay time was reduced to approximately 0.050 second, indicating a marked effect of chamber temperature at constant propellant feed temperatures. Accumulated run time affected the starting characteristics only when both the chamber and propellant feed temperatures were at reduced levels.

  4. Chemistry of the system: Al2O3(c)minus HCL aqueous. [chemical reactions resulting from propellant combustion of rocket propellants

    NASA Technical Reports Server (NTRS)

    Tyree, S. Y., Jr.

    1975-01-01

    In order to study exhaust gas chemistry for the space shuttle, the vapor pressure of 2 to 1 weight mixtures of 3-M hydrochloric acid and Al2O3 was studied over a l80 minute reaction period at 31 C. The Al2O3 sample was one of high surface area furnished by NASA Langley Research Center. A brief review is given for aqueous aluminum chemistry, and the chemical reactions of combustion products (exhaust gases) of aluminum propellant binders for the space shuttle are listed.

  5. 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.

  6. Characterization of rocket propellant combustion products. Chemical characterization and computer modeling of the exhaust products from four propellant formulations: Final report, September 23, 1987--April 1, 1990

    SciTech Connect

    Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.

    1991-12-09

    The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army`s Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

  7. Characterization of the non axial thrust generated by large solid propellant rocket motors in three axis stabilized ascent

    NASA Technical Reports Server (NTRS)

    Kosmann, W. J.; Dionne, E. R.; Klemetson, R. W.

    1978-01-01

    Nonaxial thrusts produced by solid rocket motors during three-axis stabilized attitude control have been determined from ascent experience on twenty three Burner II, Burner IIA and Block 5D-1 upper stage vehicles. A data base representing four different rocket motor designs (three spherical and one extended spherical) totaling twenty five three-axis stabilized firings is generated. Solid rocket motor time-varying resultant and lateral side force vector magnitudes, directions and total impulses, and roll torque couple magnitudes, directions, and total impulses are tabulated in the appendix. Population means and three sigma deviations are plotted. Existing applicable ground test side force and roll torque magnitudes and total impulses are evaluated and compared to the above experience data base. Within the spherical motor population, the selected AEDC ground test data consistently underestimated experienced motor side forces, roll torques and total impulses. Within the extended spherical motor population, the selected AEDC test data predicted experienced motor side forces, roll torques, and total impulses, with surprising accuracy considering the very small size of the test and experience populations.

  8. Effect of combustion-chamber pressure and nozzle expansion ratio on theoretical performance of several rocket propellant systems

    NASA Technical Reports Server (NTRS)

    Morrell, Virginia E

    1956-01-01

    Theoretical calculations of specific impulse to determine the separate effects of increasing the combustion-chamber pressure and the nozzle expansion ratio on the performance of the propellants, hydrogen-fluorine, hydrogen-oxygen, ammonia-fluorine and AN-F-58 fuel - white fuming nitric acid (95 percent). The results indicate that an increase in specific impulse obtainable with an increase in combustion-chamber pressure is almost entirely caused by the increased expansion ratio through the nozzle.

  9. Propellant variability assessment

    NASA Technical Reports Server (NTRS)

    Tytula, Thomas P.; Schad, Kristin

    1991-01-01

    Efforts to determine whether rocket propellant density and modulus can be reliably measured using non-destructive ultrasonic techniques are reported. The objective was not achieved, primarily due to the approach taken.

  10. 3-D Flash Lidar Performance in Flight Testing on the Morpheus Autonomous, Rocket-Propelled Lander to a Lunar-Like Hazard Field

    NASA Technical Reports Server (NTRS)

    Roback, Vincent E.; Amzajerdian, Farzin; Bulyshev, Alexander E.; Brewster, Paul F.; Barnes, Bruce W.

    2016-01-01

    For the first time, a 3-D imaging Flash Lidar instrument has been used in flight to scan a lunar-like hazard field, build a 3-D Digital Elevation Map (DEM), identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The flight tests served as the TRL 6 demo of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) system and included launch from NASA-Kennedy, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range. The ALHAT project developed a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The Flash Lidar is a second generation, compact, real-time, air-cooled instrument. Based upon extensive on-ground characterization at flight ranges, the Flash Lidar was shown to be capable of imaging hazards from a slant range of 1 km with an 8 cm range precision and a range accuracy better than 35 cm, both at 1-delta. The Flash Lidar identified landing hazards as small as 30 cm from the maximum slant range which Morpheus could achieve (450 m); however, under certain wind conditions it was susceptible to scintillation arising from air heated by the rocket engine and to pre-triggering on a dust cloud created during launch and transported down-range by wind.

  11. Raman Spectroscopy for Instantaneous Multipoint, Multispecies Gas Concentration and Temperature Measurements in Rocket Engine Propellant Injector Flows

    NASA Technical Reports Server (NTRS)

    Wehrmeyer, Joseph A.; Trinh, Huu Phuoc

    2001-01-01

    Propellant injector development at MSFC 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 propellant 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.

  12. Solid propellant motor

    NASA Technical Reports Server (NTRS)

    Shafer, J. I.; Marsh, H. E., Jr. (Inventor)

    1978-01-01

    A case bonded end burning solid propellant rocket motor is described. A propellant with sufficiently low modulus to avoid chamber buckling on cooling from cure and sufficiently high elongation to sustain the stresses induced without cracking is used. The propellant is zone cured within the motor case at high pressures equal to or approaching the pressure at which the motor will operate during combustion. A solid propellant motor with a burning time long enough that its spacecraft would be limited to a maximum acceleration of less than 1 g is provided by one version of the case bonded end burning solid propellant motor of the invention.

  13. Transient Simulation of Liquid Rocket Engines: A Step Towards a More Educated Propellant Choice between Kerosene and Methane

    NASA Astrophysics Data System (ADS)

    Manfletti, C.

    2004-10-01

    It is renowned that transient simulation is a powerful tool which reveals fundamental characteristics of any liquid engine. Using the program TLR-Sim (Transient Liquid Rocket Engine Simulator), whose development has recently begun, both LOX/Kerosene and LOX/Methane engine components have been closer examined whilst seeking for a clear indication of advantages associated with the use of one or the other in a number of potential engine applications. Results of the performed simulations are presented and commented. The program TLR-Sim and its development are shortly presented. A short summary of the program structure and main characteristics precedes the detailed engine analysis and comparison. A brief conclusion will follow, outlining both the main results of the inspection and a future roadmap for a subsequent program evolution.

  14. Energetic Properties of Rocket Propellants Evaluated through the Computational Determination of Heats of Formation of Nitrogen-Rich Compounds.

    PubMed

    Forquet, Valérian; Miró Sabaté, Carles; Chermette, Henry; Jacob, Guy; Labarthe, Émilie; Delalu, Henri; Darwich, Chaza

    2016-03-01

    The use of ab initio and DFT methods to calculate the enthalpies of formation of solid ionic compounds is described. The results obtained from the calculations are then compared with those from experimental measurements on nitrogen-rich salts of the 2,2-dimethyltriazanium cation (DMTZ) synthesized in our laboratory and on other nitrogen-rich ionic compounds. The importance of calculating accurate volumes and lattice enthalpies for the determination of heats of formation is also discussed. Furthermore, the crystal structure and hydrogen-bonding networks of the nitroformate salt of the DMTZ cation is described in detail. Lastly, the theoretical heats of formation were used to calculate the specific impulses (Isp ) of the salts of the DMTZ cation in view of a prospective application in propellant formulations. PMID:26762868

  15. Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

    NASA Technical Reports Server (NTRS)

    Bellman, Donald R; Humphrey, Jack C

    1948-01-01

    Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

  16. Flight Test Results of Rocket-Propelled Buffet-Research Models Having 45 Degree Sweptback Wings and 45 Degree Sweptback Tails Located in the Wing Chord Plane

    NASA Technical Reports Server (NTRS)

    Mason, Homer P.

    1953-01-01

    Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.

  17. Longitudinal Stability and Drag Characteristics at Mach Numbers from 0.70 to 1.37 of Rocket-propelled Models Having a Modified Triangular Wing

    NASA Technical Reports Server (NTRS)

    Chapman, Rowe, Jr; Morrow, John D

    1952-01-01

    A modified triangular wing of aspect ratio 2.53 having an airfoil section 3.7 percent thick at the root and 5.98 percent thick at the tip was designed in an attempt to improve the lift and drag characteristics of triangular wings. Free-flight drag and stability tests were made using rocket-propelled models equipped with the modified wing. The Mach number range of the test was from 0.70 to 1.37. Test results indicated the following: The lift-curve slope of wing plus fuselage approaches the theoretical value of wing alone at supersonic Mach numbers. The drag coefficient, based on total wing area, for wing plus interference was approximately 0.0035 at subsonic Mach numbers and 0.0080 at supersonic Mach numbers. The maximum shift in aerodynamic center for the complete configuration was 14 percent in the rearward direction from the forward position of 51.5 percent of mean aerodynamic chord at subsonic Mach numbers. The variation of lift and moment with angle of attack was linear at supersonic Mach numbers for the range of coefficients covered in the test. The high value of lift-curve slope was considered to be a significant result attributable to the wing modifications.

  18. Lidar Sensor Performance in Closed-Loop Flight Testing of the Morpheus Rocket-Propelled Lander to a Lunar-Like Hazard Field

    NASA Technical Reports Server (NTRS)

    Roback, Vincent E.; Pierrottet, Diego F.; Amzajerdian, Farzin; Barnes, Bruce W.; Hines, Glenn D.; Petway, Larry B.; Brewster, Paul F.; Kempton, Kevin S.; Bulyshev, Alexander E.

    2015-01-01

    For the first time, a suite of three lidar sensors have been used in flight to scan a lunar-like hazard field, identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, guide the Morpheus autonomous, rocket-propelled, free-flying test bed to a safe landing on the hazard field. The lidar sensors and GN&C system are part of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) project which has been seeking to develop a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The 3-D imaging flash lidar is a second generation, compact, real-time, air-cooled instrument developed from a number of cutting-edge components from industry and NASA and is used as part of the ALHAT Hazard Detection System (HDS) to scan the hazard field and build a 3-D Digital Elevation Map (DEM) in near-real time for identifying safe sites. The flash lidar is capable of identifying a 30 cm hazard from a slant range of 1 km with its 8 cm range precision at 1 sigma. The flash lidar is also used in Hazard Relative Navigation (HRN) to provide position updates down to a 250m slant range to the ALHAT navigation filter as it guides Morpheus to the safe site. The Doppler Lidar system has been developed within NASA to provide velocity measurements with an accuracy of 0.2 cm/sec and range measurements with an accuracy of 17 cm both from a maximum range of 2,200 m to a minimum range of several meters above the ground. The Doppler Lidar's measurements are fed into the ALHAT navigation filter to provide lander guidance to the safe site. The Laser Altimeter, also developed within NASA, provides range measurements with an accuracy of 5 cm from a maximum operational range of 30 km down to 1 m and, being a separate sensor from the flash lidar, can provide range along a separate vector. The Laser Altimeter measurements are also

  19. Solid Propellant Grain Structural Integrity Analysis

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The structural properties of solid propellant rocket grains were studied to determine the propellant resistance to stresses. Grain geometry, thermal properties, mechanical properties, and failure modes are discussed along with design criteria and recommended practices.

  20. Viscoelastic propellant effects on Space Shuttle Dynamics

    NASA Technical Reports Server (NTRS)

    Bugg, F.

    1981-01-01

    The program of solid propellant research performed in support of the space shuttle dynamics modeling effort is described. Stiffness, damping, and compressibility of the propellant and the effects of many variables on these properties are discussed. The relationship between the propellant and solid rocket booster dynamics during liftoff and boost flight conditions and the effects of booster vibration and propellant stiffness on free free solid rocket booster modes are described. Coupled modes of the shuttle system and the effect of propellant stiffness on the interfaces of the booster and the external tank are described. A finite shell model of the solid rocket booster was developed.

  1. 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.

  2. Lidar Sensor Performance in Closed-Loop Flight Testing of the Morpheus Rocket-Propelled Lander to a Lunar-Like Hazard Field

    NASA Technical Reports Server (NTRS)

    Roback, V. Eric; Pierrottet, Diego F.; Amzajerdian, Farzin; Barnes, Bruce W.; Bulyshev, Alexander E.; Hines, Glenn D.; Petway, Larry B.; Brewster, Paul F.; Kempton, Kevin S.

    2015-01-01

    For the first time, a suite of three lidar sensors have been used in flight to scan a lunar-like hazard field, identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The lidar sensors and GN&C system are part of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) project which has been seeking to develop a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The 3-D imaging Flash Lidar is a second generation, compact, real-time, aircooled instrument developed from a number of components from industry and NASA and is used as part of the ALHAT Hazard Detection System (HDS) to scan the hazard field and build a 3-D Digital Elevation Map (DEM) in near-real time for identifying safe sites. The Flash Lidar is capable of identifying a 30 cm hazard from a slant range of 1 km with its 8 cm range precision (1-s). The Flash Lidar is also used in Hazard Relative Navigation (HRN) to provide position updates down to a 250m slant range to the ALHAT navigation filter as it guides Morpheus to the safe site. The Navigation Doppler Lidar (NDL) system has been developed within NASA to provide velocity measurements with an accuracy of 0.2 cm/sec and range measurements with an accuracy of 17 cm both from a maximum range of 2,200 m to a minimum range of several meters above the ground. The NDLâ€"TM"s measurements are fed into the ALHAT navigation filter to provide lander guidance to the safe site. The Laser Altimeter (LA), also developed within NASA, provides range measurements with an accuracy of 5 cm from a maximum operational range of 30 km down to 1 m and, being a separate sensor from the Flash Lidar, can provide range along a separate vector. The LA measurements are also fed

  3. Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs

    NASA Technical Reports Server (NTRS)

    Keba, John E.

    1999-01-01

    This paper presents viewgraphs of The Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs. The topics include: 1) Rocket Turbomachinery Shaft Seals (Inter-Propellant-Seal (IPS) Systems, Lift-off Seal Systems, and Technology Development Needs); 2) Rocket Engine Characteristics (Engine cycles, propellants, missions, etc., Influence on shaft sealing requirements); and 3) Conclusions.

  4. 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.

  5. F. Gomez Arias' rocket vehicle project

    NASA Technical Reports Server (NTRS)

    Carreras, R.

    1977-01-01

    Research done by Spanish pioneer rocket scientists in the 19th century was investigated with major emphasis placed on F. Gomez Arias' rocket vehicle project. Arias, considered the world's first designer of rocket propelled, manned aircraft, was interested in solving the problem of space navigation. Major concerns included ascent and direction of heavier-than-airmachines, as well as ascent and direction of balloons.

  6. The Swedish Rocket Corps, 1833 - 1845

    NASA Technical Reports Server (NTRS)

    Skoog, A. I.

    1977-01-01

    Rockets for pyrotechnic displays used in Sweden in the 19th century are examined in terms of their use in war situations. Work done by the Swedish chemist J. J. Berzelius, who analyzed and improved the propellants of such rockets, and the German engineer, Martin Westermaijer, who researched manufacturing techniques of these rockets is also included.

  7. Silicone containing solid propellant

    NASA Technical Reports Server (NTRS)

    Ramohalli, K. N. R. (Inventor)

    1980-01-01

    The addition of a small amount, for example 1% by weight, of a liquid silicone oil to a metal containing solid rocket propellant provides a significant reduction in heat transfer to the inert nozzle walls. Metal oxide slag collection and blockage of the nozzle are eliminated and the burning rate is increased by about 5% to 10% thus improving ballistic performance.

  8. Longitudinal Stability Characteristics of the Consolidated Vultee XFY-1 Airplane with Windmilling Propellers as Obtained from Flight of 0.133-Scale Rocket-Propelled Model at Mach Numbers from 0.70 to 1.13

    NASA Technical Reports Server (NTRS)

    Hastings, Earl C.; Mitcham, Grady L.

    1954-01-01

    A flight test has been conducted to determine the longitudinal stability and control characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane with windmilling propellers for the Mach number range between 0.70 and 1.13. The variation of lift-curve slope C(sub L(sub alpha) with Mach number was gradual with a maximum value of 0.074 occurring at a Mach number of 0.97. Propellers had little effect upon the values of lift-curve slope or the linearity of lift coefficient with angle of attack. At lift coefficients between approximately 0.25 and 0.45 with an elevon angle of approximately -l0 deg, there was a region of neutral longitudinal stability at Mach numbers below 0.93 introduced by the addition of windmilling propellers. Below a lift coefficient of 0.10 and above a lift coefficient of 0.45, the model was longitudinally stable throughout the Mach number range of the test. There was a forward shift in the aerodynamic center of about 3-percent mean aerodynamic chord introduced by the addition of propellers. The aerodynamic center as determined at low lift moved gradually from a value of 28.5-percent mean aerodynamic chord at a Mach number of 0.75 to a value of 47-percent mean aerodynamic chord at a Mach number of 1.10. There was an abrupt decrease in pitch damping between Mach numbers of 0.88 and 0.99 followed by a rapid increase in damping to a Mach number of 1.06. The propellers had little effect upon the pitch damping characteristics . The transonic trim change was a large pitching-down tendency with and without windmilling propellers. The elevons were effective pitch controls throughout the speed range; however, their effectiveness was reduced about 50 percent at supersonic speeds. The propellers had no appreciable effect upon the control effectiveness.

  9. 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.

  10. 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.

  11. Process for the leaching of AP from propellant

    NASA Technical Reports Server (NTRS)

    Shaw, G. C.; Mcintosh, M. J. (Inventor)

    1980-01-01

    A method for the recovery of ammonium perchlorate from waste solid rocket propellant is described wherein shredded particles of the propellant are leached with an aqueous leach solution containing a low concentration of surface active agent while stirring the suspension.

  12. Summary of the Aerodynamic Characteristics and Flying Qualities Obtained from Flights of Rocket-Propelled Models of an Airplane Configuration Incorporating a Sweptback Inversely Tapered Wing at Transonic and Low-Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitcham, Grady L.; Blanchard, Willard S., Jr.

    1950-01-01

    Flight tests have been conducted on rocket-propelled models of an airplane configuration incorporating a sweptback wing with inverse taper to investigate the drag, stability, and control characteristics at transonic and supersonic speeds. The models were tested with a conventional tail arrangement in the Mach number range from 0.55 to 1.2. In addition to the various aerodynamic parameters obtained, the flying qualities were computed for a full-scale airplane with the center-of-gravity location at 18 percent of the mean aerodynamic chord. Also, included in this investigation are drag measurements made on relatively simple fixed-control models tested with both conventional and V-tail arrangements.

  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. Contact diagnostics of combustion products of rocket engines, their units, and systems

    NASA Astrophysics Data System (ADS)

    Ivanov, N. N.; Ivanov, A. N.

    2013-12-01

    This article is devoted to a new block-module device used in the diagnostics of condensed combustion products of rocket engines during research and development with liquid-propellant rocket engines (Glushko NPO Energomash; engines RD-171, RD-180, and RD-191) and solid-propellant rocket motors. Soot samplings from the supersonic high-temperature jet of a high-power liquid-propellant rocket engine were taken by the given device for the first time in practice for closed-exhaust lines. A large quantity of significant results was also obtained during a combustion investigation of solid propellants within solid-propellant rocket motors.

  15. Low acid producing solid propellants

    NASA Technical Reports Server (NTRS)

    Bennett, Robert R.

    1995-01-01

    The potential environmental effects of the exhaust products of conventional rocket propellants have been assessed by various groups. Areas of concern have included stratospheric ozone, acid rain, toxicity, air quality and global warming. Some of the studies which have been performed on this subject have concluded that while the impacts of rocket use are extremely small, there are propellant development options which have the potential to reduce those impacts even further. This paper discusses the various solid propellant options which have been proposed as being more environmentally benign than current systems by reducing HCI emissions. These options include acid neutralized, acid scavenged, and nonchlorine propellants. An assessment of the acid reducing potential and the viability of each of these options is made, based on current information. Such an assessment is needed in order to judge whether the potential improvements justify the expenditures of developing the new propellant systems.

  16. Efficiency of the rocket engines with a supersonic afterburner

    NASA Astrophysics Data System (ADS)

    Sergienko, A. A.

    1992-08-01

    The paper is concerned with the problem of regenerative cooling of the liquid-propellant rocket engine combustion chamber at high pressures of the working fluid. It is shown that high combustion product pressures can be achieved in the liquid-propellant rocket engine with a supersonic afterburner than in a liquid-propellant rocket engine with a conventional subsonic combustion chamber for the same allowable heat flux density. However, the liquid-propellant rocket engine with a supersonic afterburner becomes more economical than the conventional engine only at generator gas temperatures of 1700 K and higher.

  17. Composite propellant combustion modeling studies

    NASA Technical Reports Server (NTRS)

    Ramohalli, K.

    1977-01-01

    A review is presented of theoretical and experimental studies of composite propellant combustion. The theoretical investigations include a model of the combustion of a nonmetallized ammonium perchlorate (AP) propellant (noting time scales for vapor-phase combustion and the condensed phase) and response functions in pressure-coupled oscillations. The experimental studies are discussed with reference to scale-modeling apparatus, flame standoff distance versus velocity as a function of pressure, and results from T-burner firings of a nonmetallized AP/polysulfide propellant. Research applications including problems with nitramine propellants, the feasibility of stop-restart rockets with salt quench, and combustion problems in large boosters are outlined.

  18. Aerodynamic Characteristics and Flying Qualities of a Tailless Triangular-wing Airplane Configuration as Obtained from Flights of Rocket-propelled Models at Transonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitcham, Grady L; Stevens, Joseph E; Norris, Harry P

    1956-01-01

    A flight investigation of rocket-powered models of a tailless triangular-wing airplane configuration was made through the transonic and low supersonic speed range at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. An analysis of the aerodynamic coefficients, stability derivatives, and flying qualities based on the results obtained from the successful flight tests of three models is presented.

  19. 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.

  20. 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.

  1. 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.

  2. The Influence of the Von Opel-Valier Experiments Upon German Rocket-Propelled Model Aircraft Development, 1920's-1930's

    NASA Astrophysics Data System (ADS)

    Winter, F. H.

    The following is a brief paper presented at the XXVIIIth International Astronautical Federation (IAF) Congress held in Prague, 25 September - 1 October 1977. Although it is generally assumed that recreational model rocketry, including commercially available kits, evolved during the Space Age, there were some interesting early predecessors, especially in Germany, that apparently arose as a result of the widespread reported rocket car and plane stunts of Fritz von Opel and Max Valier in the late 1920's.

  3. 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.

  4. A Flight Demonstration of Plasma Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Petro, Andrew

    1999-01-01

    The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

  5. The scaling of the threshold conditions for solid propellant erosive burning

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Nguyen, M. H.; Cohen, N. S.

    1988-01-01

    Rocket test firings were performed to measure the transition length threshold conditions while systematically varying various rocket motor parameters. These include the crossflow velocity, the chamber pressure, the propellant nonerosive burning rate, the propellant surface roughness, and the motor port diameter. The erosive burning trends with varying propellant burning rate, motor chamber pressure, and mass flow rate are consistent with published results.

  6. Solid rocket motor internal insulation

    NASA Technical Reports Server (NTRS)

    Twichell, S. E. (Editor); Keller, R. B., Jr.

    1976-01-01

    Internal insulation in a solid rocket motor is defined as a layer of heat barrier material placed between the internal surface of the case propellant. The primary purpose is to prevent the case from reaching temperatures that endanger its structural integrity. Secondary functions of the insulation are listed and guidelines for avoiding critical problems in the development of internal insulation for rocket motors are presented.

  7. 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.

  8. Study of solid rocket motor for space shuttle booster, volume 2, book 2

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A technical analysis of the solid propellant rocket engines for use with the space shuttle is presented. The subjects discussed are: (1) solid rocket motor stage recovery, (2) environmental effects, (3) man rating of the solid propellant rocket engines, (4) system safety analysis, (5) ground support equipment, and (6) transportation, assembly, and checkout.

  9. A technology data base for the design of 500 to 5000-lb thrust class liquid rocket engines utilizing hydrogen and oxygen as propellants

    NASA Technical Reports Server (NTRS)

    Schoenman, L.

    1982-01-01

    This paper presents an overview of the results of experimental evaluations of candidate designs for igniters, injectors, and propellant-cooled thrust chambers applicable to restartable high-performance, high-reliability upper-stage engines and to pulsing-type reaction control engines (RCE). Injection element types best suited for liquid, gas, and liquid/gas phase propellant supply are identified. The resulting interactions between element type, combustion efficiency, and chamber wall heating are compared. The distinction between thrust chamber design requirements for upper stage vs RCE applications as measured by cycle life requirements is translated into design configurations consisting of all-film-cooled, all-regeneratively-cooled, and composites of the two cooling approaches. The validity of the design approaches is confirmed by data from engine durability testing involving over 90,000 starts and 9,000 thermal cycles on RCE-type designs and multiple long-duration burns (up to 2,000 sec) on regeneratively cooled upper-stage designs.

  10. 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.

  11. 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.

  12. 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.

  13. Materials characterization of propellants using ultrasonics

    NASA Technical Reports Server (NTRS)

    Workman, Gary L.; Jones, David

    1993-01-01

    Propellant characteristics for solid rocket motors were not completely determined for its use as a processing variable in today's production facilities. A major effort to determine propellant characteristics obtainable through ultrasonic measurement techniques was performed in this task. The information obtained was then used to determine the uniformity of manufacturing methods and/or the ability to determine non-uniformity in processes.

  14. Composite Solid Propellant Predictability and Quality Assurance

    NASA Technical Reports Server (NTRS)

    Ramohalli, Kumar

    1989-01-01

    Reports are presented at the meeting at the University of Arizona on the study of predictable and reliable solid rocket motors. The following subject areas were covered: present state and trends in the research of solid propellants; the University of Arizona program in solid propellants, particularly in mixing (experimental and analytical results are presented).

  15. Recovery of aluminum from composite propellants

    NASA Technical Reports Server (NTRS)

    Shaw, G. C. (Inventor)

    1980-01-01

    Aluminum was recovered from solid rocket propellant containing a small amount of oxidizer by depolymerizing and dissolving propellant binders (containing functional or hydrolyzable groups in a solution of sodium methoxide) in an alcohol solvent optionally containing an aliphatic or aromatic hydrocarbon co-solvent. The solution was filtered to recover substantially all the aluminum in active form.

  16. A smoke producing rocket motor for atmospheric wind profiling

    SciTech Connect

    Grubelich, M.C. ); Rowland, J. . Applied Physics Lab.)

    1991-01-01

    A composite propellant was developed to produce a dense plume from a rocket motor. The development of this propellant combined the smoke producing capabilities of a smoke generator with a rocket motor, thereby integrating the separate systems into one unit. A rocket motor was designed for use with this propellant to produce a high density particulate plume. This plume could then be used to determine the wind profile in the atmosphere by using a light detection and ranging system. Additionally, this smoke producing propellant could be used for rapid screening or identification. The burn rate characteristics of the propellant were measured and static firings of rocket motors were conducted to determine the performance of the propellant. The results of these tests will be presented as well as theoretical performance predictions of a flight vehicle.

  17. A smoke producing rocket motor for atmospheric wind profiling

    SciTech Connect

    Grubelich, M.C.; Rowland, J.

    1991-12-31

    A composite propellant was developed to produce a dense plume from a rocket motor. The development of this propellant combined the smoke producing capabilities of a smoke generator with a rocket motor, thereby integrating the separate systems into one unit. A rocket motor was designed for use with this propellant to produce a high density particulate plume. This plume could then be used to determine the wind profile in the atmosphere by using a light detection and ranging system. Additionally, this smoke producing propellant could be used for rapid screening or identification. The burn rate characteristics of the propellant were measured and static firings of rocket motors were conducted to determine the performance of the propellant. The results of these tests will be presented as well as theoretical performance predictions of a flight vehicle.

  18. Free-Flight Investigation of a Rocket-Propelled Model to Determine the Aerodynamic Heating on a Thin, Unswept, Untapered, Multispar, Aluminum-Alloy Wing at Mach Numbers up to 2.22

    NASA Technical Reports Server (NTRS)

    Stephens, Emily W.

    1959-01-01

    A free-flight investigation has been made to determine some effects of aerodynamic heating on the structural behavior of a wing at supersonic speeds. The test wing was a thin, unswept, untapered, multispar, aluminum-alloy wing having a 20-inch chord, a 20-inch exposed semispan, and a circular-arc airfoil section with a thickness ratio of 5 percent. The wing was tested on a model propelled by a two-stage rocket-propulsion system to a Mach number of 2.22 and a corresponding Reynolds number per foot of 13.2 x 10(6) Reasonably good agreement was obtained between Stanton numbers obtained from measured temperature-time data and values obtained by the theory of Van Driest for flat plates having turbulent boundary layers. Temperature measurements made in the skin of the wing and in the internal structures agreed well with calculated values. The wing was instrumented to detect any apparent fluttering motion in the wing, but no evidence of flutter was observed throughout the flight.

  19. Device and process for attachment of parts to rocket motors

    NASA Astrophysics Data System (ADS)

    Yagla, Jon J.; Lowry, Robert W.; Mears, Otho L.

    1992-04-01

    An attachment platform positioned longitudinally on a rocket motor chamber and secured with laser welding techniques is described. Each attachment platform is continuously sealed longitudinally to the rocket motor chamber through the application of laser welding and optical seam tracking. Application of laser welding techniques allows for repair and installation of attachment platforms on rocket motors fully loaded with live propellant.

  20. 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.

  1. Space aging of solid rocket materials

    NASA Technical Reports Server (NTRS)

    Lester, Dean M.; Jones, Leon L.; Smalley, R. B., Jr.; Ord, R. Neil

    1991-01-01

    Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. The results of aging IPSM-II/PAM-DII space motor components are presented. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results are compared to data collected in a ground based vacuum aging program on similar components.

  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. Study of solid rocket motor for space shuttle booster, volume 2, book 1

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.

  4. 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.

  5. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1987-03-01

    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report.

  6. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1986-01-01

    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report. 57 refs., 10 figs.

  7. 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

  8. 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

  9. Propellant Technologies: A Persuasive Wave of Future Propulsion Benefits

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan; Ianovski, Leonid S.; Carrick, Patrick

    1997-01-01

    Rocket propellant and propulsion technology improvements can be used to reduce the development time and operational costs of new space vehicle programs. Advanced propellant technologies can make the space vehicles safer, more operable, and higher performing. Five technology areas are described: Monopropellants, Alternative Hydrocarbons, Gelled Hydrogen, Metallized Gelled Propellants, and High Energy Density Materials. These propellants' benefits for future vehicles are outlined using mission study results and the technologies are briefly discussed.

  10. 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.

  11. Rocket Flight.

    ERIC Educational Resources Information Center

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

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

  12. 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.

  13. Refueling with In-Situ Produced Propellants

    NASA Technical Reports Server (NTRS)

    Chato, David J.

    2014-01-01

    In-situ produced propellants have been identified in many architecture studies as key to implementing feasible chemical propulsion missions to destinations beyond lunar orbit. Some of the more noteworthy ones include: launching from Mars to return to Earth (either direct from the surface, or via an orbital rendezvous); using the Earth-Moon Lagrange point as a place to refuel Mars transfer stages with Lunar surface produced propellants; and using Mars Moon Phobos as a place to produce propellants for descent and ascent stages bound for the Mars surface. However successful implementation of these strategies require an ability to successfully transfer propellants from the in-situ production equipment into the propellant tankage of the rocket stage used to move to the desired location. In many circumstances the most desirable location for this transfer to occur is in the low-gravity environment of space. In support of low earth orbit propellant depot concepts, extensive studies have been conducted on transferring propellants in-space. Most of these propellant transfer techniques will be applicable to low gravity operations in other locations. Even ground-based transfer operations on the Moon, Mars, and especially Phobos could benefit from the propellant conserving techniques used for depot refueling. This paper will review the literature of in-situ propellants and refueling to: assess the performance benefits of the use in-situ propellants for mission concepts; review the parallels with propellant depot efforts; assess the progress of the techniques required; and provide recommendations for future research.

  14. Advanced Solid Rocket Launcher and Its Evolution

    NASA Astrophysics Data System (ADS)

    Morita, Yasuhiro; Imoto, Takayuki; Habu, Hiroto; Ohtsuka, Hirohito; Hori, Keiichi; Koreki, Takemasa; Fukuchi, Apollo; Uekusa, Yasuyuki; Akiba, Ryojiro

    The research on next generation solid propellant rockets is actively underway in various spectra. JAXA is developing the Advanced Solid Rocket (ASR) as a successor to the M-V launch vehicle, which was utilized over past ten years for space science programs including planetary missions. ASR is a result of the development of the next generation technology including a highly intelligent autonomous check-out system, which is connected to not only the solid rocket but also future transportation systems. It is expected to improve the efficiency of the launch system and double the cost performance. Far beyond this effort, the passion of the volunteers among the industry-government-academia cooperation has been united to establish the society of the freewheeling thinking “Next generation Solid Rocket Society (NSRS)”. It aims at a larger revolution than what the ASR provides so that the order of the cost performance is further improved. A study of the Low melting temperature Thermoplastic Propellant (LTP) is now at the experimental stage, which is expected to reform the manufacturing process of the solid rocket propellant and lead to a significant increase in cost performance. This paper indicates the direction of the big flow towards the next generation solid-propellant rockets: the concept of the intelligent ASR under development; and the innovation behind LTP.

  15. Liquid Rocket Booster (LRB) for the Space Transportion System (STS) systems study. Appendix D: Trade study summary for the liquid rocket booster

    NASA Technical Reports Server (NTRS)

    1989-01-01

    Trade studies plans for a number of elements in the Liquid Rocket Booster (LRB) component of the Space Transportation System (STS) are given in viewgraph form. Some of the elements covered include: avionics/flight control; avionics architecture; thrust vector control studies; engine control electronics; liquid rocket propellants; propellant pressurization systems; recoverable spacecraft; cryogenic tanks; and spacecraft construction materials.

  16. 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.

  17. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

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

  18. 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)

  19. AXISYMMETRIC, THROTTLEABLE NON-GIMBALLED ROCKET ENGINE

    NASA Technical Reports Server (NTRS)

    Sackheim, Robert L. (Inventor); Hutt, John J. (Inventor); Anderson, William E. (Inventor); Dressler, Gordon A. (Inventor)

    2005-01-01

    A rocket engine assembly is provided for a vertically launched rocket vehicle. A rocket engine housing of the assembly includes two or more combustion chambers each including an outlet end defining a sonic throat area. A propellant supply for the combustion chambers includes a throttling injector, associated with each of the combustion chambers and located opposite to sonic throat area, which injects the propellant into the associated combustion chamber. A modulator, which may form part of the injector, and which is controlled by a controller, modulates the flow rate of the propellant to the combustion chambers so that the chambers provide a vectorable net thrust. An expansion nozzle or body located downstream of the throat area provides expansion of the combustion gases produced by the combustion chambers so as to increase the net thrust.

  20. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1992-01-01

    The paper describes the Advanced Solid Rocket Motor (ASRM) that is being developed to replace, in 1997, the Redesigned Solid Rocket Motor which currently boosts the Space Shuttle. The ASRM will contain features to improve motor safety (fewer potential leak paths, improved seal materials, stronger case material, and fewer nozzle and case joints), an improved ignition system using through-bulkhead initiators, and highly reproducible manufacturing and inspection techniques with a large number of automated procedures. The ASRM will be able to deliver 12,000 lbs greater payloads to any given orbit of the Shuttle. There are also environmental improvements, realized by waste propellant recovery.

  1. 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

  2. Processing solid propellants for recycling

    SciTech Connect

    Whinnery, L.L.; Griffiths, S.K.; Handrock, J.L.; Lipkin, J.

    1994-05-01

    Rapid evolution in the structure of military forces worldwide is resulting in the retirement of numerous weapon systems. Many of these systems include rocket motors containing highly energetic propellants based on hazardous nitrocellulose/nitroglycerin (NC/NG) mixtures. Even as the surplus quantities of such material increases, however, current disposal methods -- principally open burning and open detonation (OB/OD) -- are coming under close scrutiny from environmental regulators. Environmentally conscious alternatives to disposal of propellant and explosives are thus receiving renewed interest. Recycle and reuse alternatives to OB/OD appear particularly attractive because some of the energetic materials in the inventories of surplus weapon systems represent potentially valuable resources to the commercial explosives and chemical industries. The ability to reclaim such resources is therefore likely to be a key requirement of any successful technology of the future in rocket motor demilitarization. This document consists of view graphs from the poster session.

  3. Coaxial Propellant Injectors With Faceplate Annulus Control

    NASA Technical Reports Server (NTRS)

    Horn, Mark D.; Miyata, Shinjiro; Farhangi, Shahram

    2010-01-01

    An improved design concept for coaxial propellant injectors for a rocket engine (or perhaps for a non-rocket combustion chamber) offers advantages of greater robustness, less complexity, fewer parts, lower cost, and less bulk, relative to prior injectors of equivalent functionality. This design concept is particularly well suited to small, tight-tolerance injectors, for which prior designs are not suitable because the practical implementation of those designs entails very high costs and difficulty in adhering to the tolerances.

  4. Alternate propellant program, phase 1

    NASA Technical Reports Server (NTRS)

    Anderson, F. A.; West, W. R.

    1979-01-01

    Candidate propellant systems for the shuttle booster solid rocket motor (SRM), which would eliminate, or greatly reduce, the amount of HCl produced in the exhaust of the shuttle SRM were investigated. Ammonium nitrate was selected for consideration as the main oxidizer, with ammonium perchlorate and the nitramine, cyclo-tetramethylene-tetranitramine as secondary oxidizers. The amount of ammonium perchlorate used was limited to an amount which would produce an exhaust containing no more than 3% HCl.

  5. Liquid Rocket Lines, Bellows, Flexible Hoses, and Filters

    NASA Technical Reports Server (NTRS)

    1977-01-01

    Fluid-flow components in a liquid propellant rocket engine and the rocket vehicle which it propels are interconnected by lines, bellows, and flexible hoses. Elements involved in the successful design of these components are identified and current technologies pertaining to these elements are reviewed, assessed, and summarized to provide a technology base for a checklist of rules to be followed by project managers in guiding a design or assessing its adequacy. Recommended procedures for satisfying each of the design criteria are included.

  6. Lead-Free Propellant for Propellant Actuated Devices

    NASA Technical Reports Server (NTRS)

    Goodwin, John L.

    2000-01-01

    Naval Surface Warfare Center, Indian Head Division's CAD/PAD Department has been working to remove toxic compounds from our products for about a decade. In 1992, we embarked on an effort to develop a lead-free double base propellant to replace that of a foreign sole source. At the time there were availability concerns. In 1995, the department developed a strategic proposal to include a wider range of products. Efforts included such efforts as removing lead sheathing from linear explosives and replacing lead azide and lead styphnate compounds. This paper will discuss efforts specifically related to developing non-leaded double base propellant for use in various Propellant Actuated Devices (PADs) for aircrew escape systems. The propellants can replace their leaded counterparts, mitigating lead handling, processing, or toxic exposure to the environment and personnel. This work eliminates the use of leaded compounds, replacing them with a more environmentally benign metal-organic salt. Historically double-base propellants have held an advantage over other families of energetic materials through their relative insensitivity of the burning rate to changes in temperature and pressure. This desirable ballistic effect has been obtained with the use of a lead-organic salt alone or in a physical mixture with a copper-organic salt, or more recently with a lead-copper complex. These ballistic modifiers are typically added to the double-base 'paste' prior to gelatinization on heated calendars or one type or another. The effect of constant burning rate over a pressure range is called a 'plateau' while an even more beneficial effect of decreasing burning rate with increasing pressure is termed a 'mesa.' The latter effect results in very low temperature sensitivity of the propellant burning rate. Propellants with such effects are ideal tactical rocket motor propellants. The use of lead compounds poses a concern for the environment and personnel safety due to the metal's toxic

  7. Biodegradation of rocket propellent waste, ammonium perchlorate

    NASA Technical Reports Server (NTRS)

    Naqui, S. M. Z.

    1975-01-01

    The impact of the biodegradation rate of ammonium perchlorate on the environment was studied in terms of growth, metabolic rate, and total biomass of selected animal and plant species. Brief methodology and detailed results are presented.

  8. Biodegradation of rocket propellant waste, ammonium perchlorate

    NASA Technical Reports Server (NTRS)

    Naqvi, S. M. Z.; Latif, A.

    1975-01-01

    The short term effects of ammonium perchlorate on selected organisms were studied. A long term experiment was also designed to assess the changes incurred by ammonium perchlorate on the nitrogen and chloride contents of soil within a period of 3 years. In addition, an attempt was made to produce methane gas from anaerobic fermentation of the aquatic weed, Alternanthera philoxeroides.

  9. New Propellants and Cryofuels

    NASA Technical Reports Server (NTRS)

    Palasezski, Bryan; Sullivan, Neil S.; Hamida, Jaha; Kokshenev, V.

    2006-01-01

    The proposed research will investigate the stability and cryogenic properties of solid propellants that are critical to NASA s goal of realizing practical propellant designs for future spacecraft. We will determine the stability and thermal properties of a solid hydrogen-liquid helium stabilizer in a laboratory environment in order to design a practical propellant. In particular, we will explore methods of embedding atomic species and metallic nano-particulates in hydrogen matrices suspended in liquid helium. We will also measure the characteristic lifetimes and diffusion of atomic species in these candidate cryofuels. The most promising large-scale advance in rocket propulsion is the use of atomic propellants; most notably atomic hydrogen stabilized in cryogenic environments, and metallized-gelled liquid hydrogen (MGH) or densified gelled hydrogen (DGH). The new propellants offer very significant improvements over classic liquid oxygen/hydrogen fuels because of two factors: (1) the high energy-release, and (ii) the density increase per unit energy release. These two changes can lead to significant reduced mission costs and increased payload to orbit weight ratios. An achievable 5 to 10 percent improvement in specific impulse for the atomic propellants or MGH fuels can result in a doubling or tripling of system payloads. The high-energy atomic propellants must be stored in a stabilizing medium such as solid hydrogen to inhibit or delay their recombination into molecules. The goal of the proposed research is to determine the stability and thermal properties of the solid hydrogen-liquid helium stabilizer. Magnetic resonance techniques will be used to measure the thermal lifetimes and the diffusive motions of atomic species stored in solid hydrogen grains. The properties of metallic nano-particulates embedded in hydrogen matrices will also be studied and analyzed. Dynamic polarization techniques will be developed to enhance signal/noise ratios in order to be able to

  10. Rheology of composite solid propellants during motor casting

    NASA Technical Reports Server (NTRS)

    Rogers, C. J.; Smith, P. L.; Klager, K.

    1978-01-01

    In a study conducted to evaluate flow parameters of uncured solid composite propellants during motor casting, two motors (1.8M-lb grain wt) were cast with a PBAN propellant exhibiting good flow characteristics in a 260-in. dia solid rocket motor. Attention is given to the effects of propellant compositional and processing variables on apparent viscosity as they pertain to rheological behavior and grain defect formation during casting. It is noted that optimized flow behavior is impaired with solid propellant loading. Non-Newtonian pseudoplastic flow is observed, which is dependent upon applied shear stress and the age of the uncured propellant.

  11. Theoretical and Experimental Analysis of the Physics of Water Rockets

    ERIC Educational Resources Information Center

    Barrio-Perotti, R.; Blanco-Marigorta, E.; Fernandez-Francos, J.; Galdo-Vega, M.

    2010-01-01

    A simple rocket can be made using a plastic bottle filled with a volume of water and pressurized air. When opened, the air pressure pushes the water out of the bottle. This causes an increase in the bottle momentum so that it can be propelled to fairly long distances or heights. Water rockets are widely used as an educational activity, and several…

  12. Electrostatic Evaluation of the Propellant Handlers Ensemble

    NASA Technical Reports Server (NTRS)

    Hogue, Michael D.; Calle, Carlos I.; Buhler, Charles

    2006-01-01

    The Self-Contained Atmospheric Protective Ensemble (SCAPE) used in propellant handling at NASA's Kennedy Space Center (KSC) has recently completed a series of tests to determine its electrostatic properties of the coverall fabric used in the Propellant Handlers Ensemble (PHE). Understanding these electrostatic properties are fundamental to ensuring safe operations when working with flammable rocket propellants such as hydrazine, methyl hydrazine, and unsymmetrical dimethyl hydrazine. These tests include surface resistivity, charge decay, triboelectric charging, and flame incendivity. In this presentation, we will discuss the results of these tests on the current PHE as well as new fabrics and materials being evaluated for the next generation of PHE.

  13. Particle size reduction of propellants by cryocycling

    SciTech Connect

    Whinnery, L.; Griffiths, S.; Lipkin, J.

    1995-05-01

    Repeated exposure of a propellant to liquid nitrogen causes thermal stress gradients within the material resulting in cracking and particle size reduction. This process is termed cryocycling. The authors conducted a feasibility study, combining experiments on both inert and live propellants with three modeling approaches. These models provided optimized cycle times, predicted ultimate particle size, and allowed crack behavior to be explored. Process safety evaluations conducted separately indicated that cryocycling does not increase the sensitivity of the propellants examined. The results of this study suggest that cryocycling is a promising technology for the demilitarization of tactical rocket motors.

  14. 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.

  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. Solid rocket technology advancement for Space Tug and IUS applications. [Interim Upper Stage

    NASA Technical Reports Server (NTRS)

    Ascher, W.; Bailey, R. L.; Behm, J. W.; Gin, W.

    1975-01-01

    Two-burn restartable solid propellant rocket motors for the kick stage (auxiliary stage) of the Shuttle Tug, or Interim Upper Stage, are described, with details on features and test results of the ignition and quench (thrust termination) systems and procedures, fabrication of propellant and insulation, explosion hazards of propellants, and comparative data on present and future motor design. These rocket motor systems are designed for upper stage augmentation of launch vehicles and possible service in Shuttle-launched outer planet spacecraft.

  17. Analysis of a Nuclear Enhanced Airbreathing Rocket for Earth to Orbit Applications

    NASA Technical Reports Server (NTRS)

    Adams, Robert B.; Landrum, D. Brian; Brown, Norman (Technical Monitor)

    2001-01-01

    The proposed engine concept is the Nuclear Enhanced Airbreathing Rocket (NEAR). The NEAR concept uses a fission reactor to thermally heat a propellant in a rocket plenum. The rocket is shrouded, thus the exhaust mixes with ingested air to provide additional thermal energy through combustion. The combusted flow is then expanded through a nozzle to provide thrust.

  18. Thrust engine and propellant exhaust arrangement therefor

    SciTech Connect

    Retallick, F.D.

    1981-01-27

    A nuclear engine and nozzle arrangement are described for a nuclear rocket comprising a cluster of elongated fissile fuel bearing and high temperature capacity modules suitably supported in a pressure vessel. The modules have a plurality of coolant-propellant channels extending therethrough, a convergent - divergent nozzle structure of fixed cross-sectional dimensions secured to the end portion of each of said modules, a divergent-only unitary skirt member connected directly to the propellant exit end of said modular cluster in series with and diverging from the divergent ends of said convergent-divergent nozzle structures. The modules are formed to conduct a compressible propellant therethrough at sub-sonic velocities, said nozzle structures being formed to develop supersonic velocities of the propellant and said divergent-only skirt being formed to develop further the supersonic velocities of said propellant.

  19. SRM propellant, friction/ESD testing

    NASA Technical Reports Server (NTRS)

    Campbell, L. A.

    1989-01-01

    Following the Pershing 2 incident in 1985 and the Peacekeeper ignition during core removal in 1987, it was found that propellant can be much more sensitive to Electrostatic Discharges (ESD) than ever before realized. As a result of the Peacekeeper motor near miss incident, a friction machine was designed and fabricated, and used to determine friction hazards during core removal. Friction testing with and electrical charge being applied across the friction plates resulted in propellant ignitions at low friction pressures and extremely low ESD levels. The objective of this test series was to determine the sensitivity of solid rocket propellant to combined friction pressure and electrostatic stimuli and to compare the sensitivity of the SRM propellant to Peacekeeper propellant. The tests are fully discussed, summarized and conclusions drawn.

  20. 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

  1. 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

  2. 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)

  3. 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.

  4. 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.

  5. Nondestructive testing of brazed rocket engine components

    NASA Technical Reports Server (NTRS)

    Adams, C. J.; Hagemaier, D. J.; Meyer, J. A.

    1968-01-01

    Report details study made of nondestructive radiographic, ultrasonic, thermographic, and leak test methods used to inspect and evaluate the quality of the various brazed joints in liquid-propellant rocket engine components and assemblies. Descriptions of some of the unique equipment and methods developed are included.

  6. Remote rocket engine leak detection techniques

    NASA Astrophysics Data System (ADS)

    Maram, J. M.

    1993-06-01

    Optical imaging techniques have been successfully used for detection of leaks in rocket engine components using introduced gases such as sulfur hexafluoride and nitrous oxide. The approach used and past applications of the technology are described. The potential for direct detection of launch system propellant leaks is discussed.

  7. The development of space solid rocket motors in China

    NASA Astrophysics Data System (ADS)

    Jianding, Huang; Dingyou, Ye

    1997-01-01

    China has undertaken to research and develop composite solid propellant rocket motors since 1958. At the request of the development of space technology, composite solid propellant rocket motor has developed from small to large, step by step. For the past thirty eight years, much progress has made, many technical obstacles, such as motor design, case materials and their processing technology, propellant formulations and manufacture, nozzles and thrust vector control, safe ignition, environment tests, nondestructive inspection and quality assurance, static firing test and measurement etc. have been solved. A serial of solid rocket motors have been offered for China's satellites launch. The systems of research, design, test and manufacture of solid rocket motors have been formed.

  8. 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.

  9. 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)

  10. Study of solid rocket motor for space, shuttle booster, volume 2, book 4 appendices B thru D

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The mass properties and related data for the solid propellant rocket engine for use with the space shuttle are presented. Data for three solid propellant rocket engines are provided. The three designs considered are: (1) baseline parallel burn, (2) optional parallel burn, and (3) baseline series burn. Layouts of the respective designs to show design and dimensional data are included.

  11. Solid propellant environmental issues

    SciTech Connect

    Le, M.D.

    1998-07-01

    The objective of the Solid Propellant Environmental Issues (SPEI) project is to demonstrate environmentally acceptable technologies that will enhance the continued production of solid rocket motors (SRMs) by complying with current and anticipated environmental regulations. Phase 1 of the project identifies current and anticipated environmental regulations that may affect SRMs manufacturing in the future and identify emerging process technologies which comply with these regulations. Phase 2 of the project established a baseline database by fabricating a 363 kg motor using the current manufacturing process. In Phase 3, environmentally acceptable process technologies were evaluated, ranked, and selected for demonstration using criteria developed by the team. The results for Phase 1--3 have previously been presented. This paper will present data obtained to date on Phase 4. In Phase 4, the alternate process technologies were evaluated for compatibility, cleaning effectiveness, and waste minimization/pollution prevention. The best performing candidate for each application area was selected for demonstration. The selected process technologies will be inserted into the baseline manufacturing process from Phase 2. The new manufacturing process will be demonstrated and evaluated through the scale-up and fabrication of two 363 kg solid rocket motors.

  12. 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.

  13. 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.

  14. 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.

  15. The alleged contributions of Pedro E. Paulet to liquid-propellant rocketry

    NASA Technical Reports Server (NTRS)

    Ordway, F. I., III

    1977-01-01

    The first practical working liquid propellant rocket motor was claimed by Pedro E. Paulet, a South American engineer from Peru (1895). He operated a conical motor, 10 centimeters in diameter, using nitrogen peroxide and gasoline as propellants and measuring thrust up to 90 kilograms, and apparently used spark ignition and intermittent propellant injection. The test device which he used contained elements of later test stands, such as a spring thrust-measuring device. However, he did not publish his work until twenty-five years later. Evidence is examined concerning this only known claim to liquid propellant rocket engine experiments in the nineteenth century.

  16. Low-Cost Propellant Launch From a Tethered Balloon

    NASA Technical Reports Server (NTRS)

    Wilcox, Brian

    2006-01-01

    A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

  17. Low thrust rocket test facility

    NASA Technical Reports Server (NTRS)

    Arrington, Lynn A.; Schneider, Steven J.

    1990-01-01

    A low thrust chemical rocket test facility has recently become operational at the NASA-Lewis. The new facility is used to conduct both long duration and performance tests at altitude over a thruster's operating envelope using hydrogen and oxygen gas for propellants. The facility provides experimental support for a broad range of objectives, including fundamental modeling of fluids and combustion phenomena, the evaluation of thruster components, and life testing of full rocket designs. The major mechanical and electrical systems are described along with aspects of the various optical diagnostics available in the test cell. The electrical and mechanical systems are designed for low down time between tests and low staffing requirements for test operations. Initial results are also presented which illustrate the various capabilities of the cell.

  18. Relativistic rocket: Dream and reality

    NASA Astrophysics Data System (ADS)

    Semyonov, Oleg G.

    2014-06-01

    The dream of interstellar flights persists since the first pioneers in astronautics and has never died. Many concepts of thruster capable to propel a rocket to the stars have been proposed and the most suitable among them are thought to be photon propulsion and propulsion by the products of proton-antiproton annihilation in magnetic nozzle. This article addresses both concepts allowing for cross-section of annihilation among other issues in order to show their vulnerability and to indicate the problems. The concept of relativistic matter propulsion is substantiated and discussed. The latter is argued to be the most straightforward way to build-up a relativistic rocket firstly because it is based on the existing technology of ion generators and accelerators and secondly because it can be stepped up in efflux power starting from interplanetary spacecrafts powered by nuclear reactors to interstellar starships powered by annihilation reactors. The problems imposed by thermodynamics and heat disposal are accentuated.

  19. Space Shuttle solid rocket booster

    NASA Technical Reports Server (NTRS)

    Hardy, G. B.

    1979-01-01

    Details of the design, operation, testing and recovery procedures of the reusable solid rocket boosters (SRB) are given. Using a composite PBAN propellant, they will provide the primary thrust (six million pounds maximum at 20 s after ignition) within a 3 g acceleration constraint, as well as thrust vector control for the Space Shuttle. The drogues were tested to a load of 305,000 pounds, and the main parachutes to 205,000. Insulation in the solid rocket motor (SRM) will be provided by asbestos-silica dioxide filled acrylonitrile butadiene rubber ('asbestos filled NBR') except in high erosion areas (principally in the aft dome), where a carbon-filled ethylene propylene diene monomer-neopreme rubber will be utilized. Furthermore, twenty uses for the SRM nozzle will be allowed by its ablative materials, which are principally carbon cloth and silica cloth phenolics.

  20. 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.

  1. Solid rocket motor temperature sensitivity

    SciTech Connect

    Osborn, J.R.; Heister, S.D.

    1994-11-01

    The temperature sensitivity of the propellant and the solid rocket motor are described by several different temperature sensitivity coefficients. This enabled the derivation of three different relationships for the temperature sensitivity coefficient pi(sub K). To demonstrate this, two different propellants were used wherein the values of pi(sub K) were generated and compared. It was observed that the expressions are of equal complexity and offer ease of use. All involve only the burning rate data and the use of the parameters in St. Roberts burning rate low. It is also suggested that the most general expression for the sensitivity coefficient should be used since it is a true pi(sub K) relationship having the partial derivatives taken with the motor geometry held constant. 11 refs.

  2. The Advanced Solid Rocket Motor

    NASA Astrophysics Data System (ADS)

    Mitchell, Royce E.

    1992-08-01

    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  3. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1992-01-01

    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  4. Advances in Green Cryogenic Solid Propellant Propulsion

    NASA Astrophysics Data System (ADS)

    Lo, R. E.; Adirim, H.; Poller, S.; Glaeser, S.; Schoeyer, H.; Caramelli, F.

    2004-10-01

    The combustion of hydrocarbons with hydrogen peroxide or oxygen based oxidizers is known as the best possible realization of green bipropellants in the realm of conventional propellants. By the nature of these constituents, corresponding rocket motors are either hybrids or bi-liquids. This is advantageous in all applications requiring the merits of these categories, such as variations of the thrust - time profile (throttle-ability up to shut down and restart), or variable propellant loading and mixture ratio variation in liquid bipropellants. However, when it comes to thriving on the simplicity and reliability of solid propellant technology, it takes cryogenic solid propulsion (CSP) as enabling technology to make these normally liquid propellants available for many solid propellant applications, in particular for high thrust Earth-to-orbit boosting. It is obvious that proper CSP propellant selection yields solids that are as "green" as any chemical propellant combination can be. The paper describes recent advances in CSP technology related investigations sponsored by the German Aerospace Centre DLR and the European Space Agency ESA at AI/ICT.

  5. 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.

  6. Dynamic characterization and analysis of space shuttle SRM solid propellant

    NASA Technical Reports Server (NTRS)

    Hufferd, W. L.

    1979-01-01

    The dynamic response properties of the space shuttle solid rocket moter (TP-H1148) propellant were characterized and the expected limits of propellant variability were established. Dynamic shear modulus tests conducted on six production batches of TP-H1148 at various static and dynamic strain levels over the temperature range from 40 F to 90 F. A heat conduction analysis and dynamic response analysis of the space shuttle solid rocket motor (SRM) were also conducted. The dynamic test results show significant dependence on static and dynamic strain levels and considerable batch-to-batch and within-batch variability. However, the results of the SRM dynamic response analyses clearly demonstrate that the stiffness of the propellant has no consequential on the overall SRM dynamic response. Only the mass of the propellant needs to be considered in the dynamic analysis of the space shuttle SRM.

  7. Interactive Schematic Integration Within the Propellant System Modeling Environment

    NASA Technical Reports Server (NTRS)

    Coote, David; Ryan, Harry; Burton, Kenneth; McKinney, Lee; Woodman, Don

    2012-01-01

    Task requirements for rocket propulsion test preparations of the test stand facilities drive the need to model the test facility propellant systems prior to constructing physical modifications. The Propellant System Modeling Environment (PSME) is an initiative designed to enable increased efficiency and expanded capabilities to a broader base of NASA engineers in the use of modeling and simulation (M&S) technologies for rocket propulsion test and launch mission requirements. PSME will enable a wider scope of users to utilize M&S of propulsion test and launch facilities for predictive and post-analysis functionality by offering a clean, easy-to-use, high-performance application environment.

  8. Atomic hydrogen propellants: Historical perspectives and future possibilities

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    1993-01-01

    Atomic hydrogen, a very high density free-radical propellant, is anticipated to generate a specific impulse of 600-1500 lb-f sec/lb-mass performance; this may facilitate the development of unique launch vehicles. A development status evaluation is presently given for atomic hydrogen investigations. It is noted that breakthroughs are required in the production, storage, and transfer of atomic hydrogen, before this fuel can become a viable rocket propellant.

  9. Bistable (latching) solenoid actuated propellant isolation valve

    NASA Technical Reports Server (NTRS)

    Wichmann, H.; Deboi, H. H.

    1979-01-01

    The design, fabrication, assembly and test of a development configuration bistable (latching) solenoid actuated propellant isolation valve suitable for the control hydrazine and liquid fluorine to an 800 pound thrust rocket engine is described. The valve features a balanced poppet, utilizing metal bellows, a hard poppet/seat interface and a flexure support system for the internal moving components. This support system eliminates sliding surfaces, thereby rendering the valve free of self generated particles.

  10. Environmental Effects of Space Shuttle Solid Rocket Motor Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Hwang, B.; Pergament, H. S.

    1976-01-01

    The deposition of NOx and HCl in the stratosphere from the space shuttle solid rocket motors (SRM) and exhaust plume is discussed. A detailed comparison between stratospheric deposition rates using the baseline SRM propellant and an alternate propellant, which replaces ammonium perchlorate by ammonium nitrate, shows the total NOx deposition rate to be approximately the same for each propellant. For both propellants the ratio of the deposition rates of NOx to total chlorine-containing species is negligibly small. Rocket exhaust ground cloud transport processes in the troposphere are also examined. A brief critique of the multilayer diffusion models (presently used for predicting pollutant deposition in the troposphere) is presented, and some detailed cloud rise calculations are compared with data for Titan 3C launches. The results show that, when launch time meteorological data are used as input, the model can reasonably predict measured cloud stabilization heights.

  11. Solid rocket technology advancements for space tug and IUS applications

    NASA Technical Reports Server (NTRS)

    Ascher, W.; Bailey, R. L.; Behm, J. W.; Gin, W.

    1975-01-01

    In order for the shuttle tug or interim upper stage (IUS) to capture all the missions in the current mission model for the tug and the IUS, an auxiliary or kick stage, using a solid propellant rocket motor, is required. Two solid propellant rocket motor technology concepts are described. One concept, called the 'advanced propulsion module' motor, is an 1800-kg, high-mass-fraction motor, which is single-burn and contains Class 2 propellent. The other concept, called the high energy upper stage restartable solid, is a two-burn (stop-restartable on command) motor which at present contains 1400 kg of Class 7 propellant. The details and status of the motor design and component and motor test results to date are presented, along with the schedule for future work.

  12. Analytic Modeling of Pressurization and Cryogenic Propellant

    NASA Technical Reports Server (NTRS)

    Corpening, Jeremy H.

    2010-01-01

    An analytic model for pressurization and cryogenic propellant conditions during all mission phases of any liquid rocket based vehicle has been developed and validated. The model assumes the propellant tanks to be divided into five nodes and also implements an empirical correlation for liquid stratification if desired. The five nodes include a tank wall node exposed to ullage gas, an ullage gas node, a saturated propellant vapor node at the liquid-vapor interface, a liquid node, and a tank wall node exposed to liquid. The conservation equations of mass and energy are then applied across all the node boundaries and, with the use of perfect gas assumptions, explicit solutions for ullage and liquid conditions are derived. All fluid properties are updated real time using NIST Refprop.1 Further, mass transfer at the liquid-vapor interface is included in the form of evaporation, bulk boiling of liquid propellant, and condensation given the appropriate conditions for each. Model validation has proven highly successful against previous analytic models and various Saturn era test data and reasonably successful against more recent LH2 tank self pressurization ground test data. Finally, this model has been applied to numerous design iterations for the Altair Lunar Lander, Ares V Core Stage, and Ares V Earth Departure Stage in order to characterize Helium and autogenous pressurant requirements, propellant lost to evaporation and thermodynamic venting to maintain propellant conditions, and non-uniform tank draining in configurations utilizing multiple LH2 or LO2 propellant tanks. In conclusion, this model provides an accurate and efficient means of analyzing multiple design configurations for any cryogenic propellant tank in launch, low-acceleration coast, or in-space maneuvering and supplies the user with pressurization requirements, unusable propellants from evaporation and liquid stratification, and general ullage gas, liquid, and tank wall conditions as functions of time.

  13. Solid Rocket Booster Structural Test Article

    NASA Technical Reports Server (NTRS)

    1978-01-01

    The structural test article to be used in the solid rocket booster (SRB) structural and load verification tests is being assembled in a high bay building of the Marshall Space Flight Center (MSFC). The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment.

  14. Ignition transient analysis of solid rocket motor

    NASA Technical Reports Server (NTRS)

    Han, Samuel S.

    1991-01-01

    Measurement data on the performance of Space Shuttle Solid Rocket Motor show wide variations in the head-end pressure changes and the total thrust build-up during the ignition transient periods. To analyze the flow and thermal behavior in the tested solid rocket motors, a 1-dimensional, ideal gas flow model via the SIMPLE algorithm was developed. Numerical results showed that burning patterns in the star-shaped head-end segment of the propellant and the erosive burning rate are two important factors controlling the ignition transients. The objective of this study is to extend the model to include the effects of aluminum particle commonly used in solid propellants. To treat the effects of aluminum-oxide particles in the combustion gas, conservation of mass, momentum, and energy equations for the particles are added in the numerical formulation and integrated by an inter-phase-slip algorithm.

  15. Solid Rocket Booster-Illustration

    NASA Technical Reports Server (NTRS)

    1977-01-01

    This illustration is a cutaway of the solid rocket booster (SRB) sections with callouts. The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment. The boosters are designed to survive water impact at almost 60 miles per hour, maintain flotation with minimal damage, and preclude corrosion of the hardware exposed to the harsh seawater environment. Under the project management of the Marshall Space Flight Center, the SRB's are assembled and refurbished by the United Space Boosters. The SRM's are provided by the Morton Thiokol Corporation.

  16. Ignition transient analysis of solid rocket motor

    NASA Astrophysics Data System (ADS)

    Han, Samuel S.

    A 1-D numerical model based on the SIMPLE is developed to predict the pressure and thrust behavior of space shuttle solid rocket motors. The present model solves the conservation equations through the attached nozzle as well as in the combustion chamber. Numerical results were seen to agree qualitatively well with the test data by controlling the wetted perimeter in the head-end star-section of the motor and the erosive burning rate of the solid propellent.

  17. Reliability Estimation Methods for Liquid Rocket Engines

    NASA Astrophysics Data System (ADS)

    Hirata, Kunio; Masuya, Goro; Kamijo, Kenjiro

    Reliability estimation using the dispersive, binominal distribution method has been traditionally used to certify the reliability of liquid rocket engines, but its estimation sometimes disagreed with the failure rates of flight engines. In order to take better results, the reliability growth model and the failure distribution method are applied to estimate the reliability of LE-7A engines, which have propelled the first stage of H-2A launch vehicles.

  18. Space shuttle propellant constitutive law verification tests

    NASA Technical Reports Server (NTRS)

    Thompson, James R.

    1995-01-01

    As part of the Propellants Task (Task 2.0) on the Solid Propulsion Integrity Program (SPIP), a database of material properties was generated for the Space Shuttle Redesigned Solid Rocket Motor (RSRM) PBAN-based propellant. A parallel effort on the Propellants Task was the generation of an improved constitutive theory for the PBAN propellant suitable for use in a finite element analysis (FEA) of the RSRM. The outcome of an analysis with the improved constitutive theory would be more reliable prediction of structural margins of safety. The work described in this report was performed by Materials Laboratory personnel at Thiokol Corporation/Huntsville Division under NASA contract NAS8-39619, Mod. 3. The report documents the test procedures for the refinement and verification tests for the improved Space Shuttle RSRM propellant material model, and summarizes the resulting test data. TP-H1148 propellant obtained from mix E660411 (manufactured February 1989) which had experienced ambient igloo storage in Huntsville, Alabama since January 1990, was used for these tests.

  19. Random sphere packing model of heterogeneous propellants

    NASA Astrophysics Data System (ADS)

    Kochevets, Sergei Victorovich

    propellants is developed which uses the concept of multi-point correlation functions. A set of intrinsic length scales of local density fluctuations in random heterogeneous propellants is identified by performing a Monte-Carlo study of the correlation functions. This method of analysis shows great promise for understanding the origins of the combustion instability of heterogeneous propellants, and is believed to become a valuable tool for the development of safe and reliable rocket engines.

  20. Propellant Management and Conditioning within the X-34 Main Propulsion System

    NASA Technical Reports Server (NTRS)

    Brown, T. M.; McDonald, J. P.; Hedayat, A.; Knight, K. C.; Champion, R. H., Jr.

    1998-01-01

    The X-34 hypersonic flight vehicle is currently under development by Orbital Sciences Corporation (Orbital). The Main Propulsion ystem as been designed around the liquid propellant Fastrac rocket engine currently under development at NASA Marshall Space Flight Center. This paper presents analyses of the MPS subsystems used to manage the liquid propellants. These subsystems include the propellant tanks, the tank vent/relief subsystem, and the dump/fill/drain subsystem. Analyses include LOX tank chill and fill time estimates, LOX boil-off estimates, propellant conditioning simulations, and transient propellant dump simulations.

  1. 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.

  2. ELIMINATION OF TOXIC MATERIALS AND SOLVENTS FROM SOLID PROPELLANT COMPONENTS (PP1058)

    EPA Science Inventory

    Solid rocket motor (srm) propellant used in military tactical and strategic missile weapon systems contain hazardous and toxic materials. Lead is used as a ballistic catalyst in minimum signature propellants. It's use results in approximately 161,000 pounds of lead compound usag...

  3. Space aging of solid rocket materials

    NASA Technical Reports Server (NTRS)

    Lester, Dean M.; Jones, Leon L.; Smalley, R. B., Jr.; Ord, R. Neil

    1992-01-01

    Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. This paper will present the results of aging the Improved Performance Space Motor-II/Payload Assist Module-Delta II (IPSM-II/PAM-DII) space motor components. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results will be compared to data collected in a ground based vacuum aging program on similar components. The operation of the vacuum actuated venting valve and pressure actuated resealing of the container will also be addressed. The materials tested showed no significant changes due to space aging. These results indicate that properly designed solid rocket motors can be expected to perform reliably after extended periods of exposure to a space environment.

  4. Overview of the manufacturing sequence of the Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Chapman, John S.; Nix, Michael B.

    1992-01-01

    The manufacturing sequence of NASA's new Advanced Solid Rocket Motor, developed as a replacement of the Space Shuttle's existing Redesigned Solid Rocket Motor, is overviewed. Special attention is given to the case preparation, the propellant mix/cast, the nondestructuve evaluation, the motor finishing, and the refurbishment. The fabrication sequences of the case, the nozzle, and the igniter are described.

  5. SRM propellant and polymer materials structural test program

    NASA Technical Reports Server (NTRS)

    Moore, Carleton J.

    1988-01-01

    The SRM propellant and polymer materials structural test program has potentially wide application to the testing and structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. The test program will provide a basis for characterization of the dynamic failure criteria for Solid Rocket Motor (SRM) propellant, insulation, inhibitor and liners. This experimental investigation will also endeavor to obtain a consistent complete set of materials test data. This test will be used to improve and revise the presently used theoretical math models for SRM propellant, insulators, inhibitor, liners, and O-ring seals.

  6. Fluid mechanics of spinning rockets

    NASA Astrophysics Data System (ADS)

    Flandro, G. A.; Vanmoorhem, W. K.; Shorthill, R.; Chen, K.; Woolsey, M.

    1987-01-01

    This report presents the results of a detailed investigation of the influence of time-dependent combustion gas flows on the attitude dynamics of spinning rocket propelled space vehicles. The work was motivated by a need to understand the origins of a potentially serious system performance problem first detected in the PAM-D series of spin stabilized upper stages. Small wobbling (often referred to as nutation or coning) is induced during separation of the rocket motor burn. The growth ceased abruptly at motor burnout, and final cone angles as large as 17 deg were reached in some flights. The same phenomenon was encountered in two flights of the PAM-DII, a similar vehicle utilizing a larger motor. Conventional theories of spinning rocket dynamics failed to explain this behavior. Since the telemetry data shows that the severity of the problem depends on spacecraft mass properties and other system parameters, it is crucial that the origins of the instability be understood completely in order that serious mission degradation can be avoided in future orbit raising operations. A costly interim fix, which sidesteps the need to understand the physical origins of the problem, is the use of a strap-on nutation control system as used in the Air Force SGS II missions.

  7. Rocket Engine Altitude Simulation Technologies

    NASA Technical Reports Server (NTRS)

    Woods, Jody L.; Lansaw, John

    2010-01-01

    John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

  8. Solar rocket system concept analysis

    NASA Technical Reports Server (NTRS)

    Boddy, J. A.

    1980-01-01

    The use of solar energy to heat propellant for application to Earth orbital/planetary propulsion systems is of interest because of its performance capabilities. The achievable specific impulse values are approximately double those delivered by a chemical rocket system, and the thrust is at least an order of magnitude greater than that produced by a mercury bombardment ion propulsion thruster. The primary advantage the solar heater thruster has over a mercury ion bombardment system is that its significantly higher thrust permits a marked reduction in mission trip time. The development of the space transportation system, offers the opportunity to utilize the full performance potential of the solar rocket. The requirements for transfer from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) was examined as the return trip, GEO to LEO, both with and without payload. Payload weights considered ranged from 2000 to 100,000 pounds. The performance of the solar rocket was compared with that provided by LO2-LH2, N2O4-MMH, and mercury ion bombardment systems.

  9. Flight Investigation of 6.25-Inch-Diameter Deacon Rocket and 10-Inch-Scale Model Rocket

    NASA Technical Reports Server (NTRS)

    Watson, R. S.

    1949-01-01

    Flight tests were conducted at the NACA Pilotless Aircraft Research Station, Wallops Island, to determine the characteristics of the Allegany Ballistics Laboratory's 6.2inch-diameter Deacon and lO-inch-scale model solid-propellant rocket motors. The tests were performed to assist in the development of these rockets which were designed for, and urgently needed to propel supersonic research models and pilotless aircraft. The tests showed that the rocket motors functioned properly under various flight- acceleration loads over a range of pre-ignition grain temperatures. A maximum velocity of 4180 feet per second was obtained at an elapsed time of 2.9 seconds with the 6.25-inch Deacon rocket motor at a gross weight of l9O pounds. Free-flight data of drag coefficient for the Deacon configuration for a Mach number range of 1.1 to 3.6 have been obtained from flight tests of several pounds. Camera studies of the take-off and flights of the Deacon rocket shared no evidence of breakup of propellant grains. An analysis of the forces to which the Deacon rocket grain is subjected was made. The analysis shows that the grain loading is most severe near the beginning and near the end of the rocket action time. The 10-inch-scale model rocket motor is a scaled model of the l6-inch- diameter multi-perforated, cast-grain rocket motor. A maximum velocity of 1625 feet per second at a time of 1.075 seconds was obtained at a gross weight of 309 pounds.

  10. Development of nuclear rocket engine technology

    SciTech Connect

    Gunn, S.V.

    1989-01-01

    Research sponsored by the Atomic Energy Commission, the USAF, and NASA (later on) in the area of nuclear rocket propulsion is discussed. It was found that a graphite reactor, loaded with highly concentrated Uranium 235, can be used to heat high pressure liquid hydrogen to temperatures of about 4500 R, and to expand the hydrogen through a high expansion ratio rocket nozzle assembly. The results of 20 reactor tests conducted at the Nevada Test Site between July 1959 and June 1969 are analyzed. On the basis of these results, the feasibility of solid graphite reactor/nuclear rocket engines is revealed. It is maintained that this technology will support future space propulsion requirements, using liquid hydrogen as the propellant, for thrust requirements ranging from 25,000 lbs to 250,000 lbs, with vacuum specific impulses of at least 850 sec and with full engine throttle capability. 12 refs.

  11. Atmospheric scavenging of solid rocket exhaust effluents

    NASA Technical Reports Server (NTRS)

    Fenton, D. L.; Purcell, R. Y.

    1978-01-01

    Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. Two chambers were used to conduct the experiments; a large, rigid walled, spherical chamber stored the exhaust constituents, while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique used. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity. Characterization of the aluminum oxide particles substantiated the similarity between the constituents of the small scale rocket and the full size vehicles.

  12. Technology for low cost solid rocket boosters.

    NASA Technical Reports Server (NTRS)

    Ciepluch, C.

    1971-01-01

    A review of low cost large solid rocket motors developed at the Lewis Research Center is given. An estimate is made of the total cost reduction obtainable by incorporating this new technology package into the rocket motor design. The propellant, case material, insulation, nozzle ablatives, and thrust vector control are discussed. The effect of the new technology on motor cost is calculated for a typical expandable 260-in. booster application. Included in the cost analysis is the influence of motor performance variations due to specific impulse and weight changes. It is found for this application that motor costs may be reduced by up to 30% and that the economic attractiveness of future large solid rocket motors will be improved when the new technology is implemented.

  13. Fast burning propellants

    SciTech Connect

    Colgate, S.A.; Roos, G.E.

    1987-07-21

    A solid or semisolid propellant is described comprising grains of propellant or propellant components bonded together to create voids within the propellant volume. The grains are of near-uniform size and have less than about a 20% size variation between the largest and smallest grains, the voids comprising from about 10% to about 50% of the propellant volume. The grains are bonded together with sufficient strength to substantially delay the fluidization of the propellant by the onset of Taylor unstable burning. The propellant has a rapid burn rate of from about 10 cm sec/sup -1/ to about 10/sup 4/cm sec/sup -1/.

  14. 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.

  15. A Flight Demonstration of Plasma Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Petro, Andrew; Chang-Diaz, Franklin; Schwenterly, WIlliam; Hitt, Michael; Lepore, Joseph

    2000-01-01

    The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a variable specific impulse magnetoplasma rocket (V ASIMR) for several years. This type of rocket could be used in the future to propel interplanetary spacecraft and has the potential to open the entire solar system to human exploration. One feature of this propulsion technology is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. Variation of specific impulse and thrust enhances the ability to optimize interplanetary trajectories and results in shorter trip times and lower propellant requirements than with a fixed specific impulse. In its ultimate application for interplanetary travel, the VASIMR would be a multi-megawatt device. A much lower power system is being designed for demonstration in the 2004 timeframe. This first space demonstration would employ a lO-kilowatt thruster aboard a solar powered spacecraft in Earth orbit. The 1O-kilowatt V ASIMR demonstration unit would operate for a period of several months with hydrogen or deuterium propellant with a specific impulse of 10,000 seconds.

  16. A review of research in low earth orbit propellant collection

    NASA Astrophysics Data System (ADS)

    Singh, Lake A.; Walker, Mitchell L. R.

    2015-05-01

    This comprehensive review examines the efforts of previous researchers to develop concepts for propellant-collecting spacecraft, estimate the performance of these systems, and understand the physics involved. Rocket propulsion requires the spacecraft to expend two fundamental quantities: energy and propellant mass. A growing number of spacecraft collect the energy they need to execute propulsive maneuvers in-situ with solar panels. In contrast, every spacecraft using rocket propulsion has carried all of the propellant mass needed for the mission from the ground, which limits the range and mission capabilities. Numerous researchers have explored the concept of collecting propellant mass while in space. These concepts have varied in scale and complexity from chemical ramjets to fusion-driven interstellar vessels. Research into propellant-collecting concepts occurred in distinct eras. During the Cold War, concepts tended to be large, complex, and nuclear powered. After the Cold War, concepts transitioned to solar power sources and more effort has been devoted to detailed analysis of specific components of the propellant-collecting architecture. By detailing the major contributions and limitations of previous work, this review concisely presents the state-of-the-art and outlines five areas for continued research. These areas include air-compatible cathode technology, techniques to improve propellant utilization on atmospheric species, in-space compressor and liquefaction technology, improved hypersonic and hyperthermal free molecular flow inlet designs, and improved understanding of how design parameters affect system performance.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 62. Historic propellant piping diagram of oxidant pit at Building ...

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

    62. Historic propellant piping diagram of oxidant pit at Building 202, January 6, 1956. NASA GRC drawing no. CF-101644. (On file at NASA Glenn Research Center). - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  3. Characterizing high-energy-density propellants for space propulsion applications

    NASA Astrophysics Data System (ADS)

    Kokan, Timothy

    There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement for existing industry standard fuels for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, the NASA Marshall Space Flight Center, and the NASA Glenn Research Center each either recently concluded or currently has ongoing programs in the synthesis and development of these potential new propellants. In order to perform conceptual designs using these new propellants, most conceptual rocket engine powerhead design tools (e.g. NPSS, ROCETS, and REDTOP-2) require several thermophysical properties of a given propellant over a wide range of temperature and pressure. These properties include enthalpy, entropy, density, viscosity, and thermal conductivity. Very little thermophysical property data exists for most of these potential new HEDM propellants. Experimental testing of these properties is both expensive and time consuming and is impractical in a conceptual vehicle design environment. A new technique for determining these thermophysical properties of potential new rocket engine propellants is presented. The technique uses a combination of three different computational methods to determine these properties. Quantum mechanics and molecular dynamics are used to model new propellants at a molecular level in order to calculate density, enthalpy, and entropy. Additivity methods are used to calculate the kinematic viscosity and thermal conductivity of new propellants. This new technique is validated via a series of verification experiments of HEDM compounds. Results are provided for two HEDM propellants: quadricyclane and 2-azido-N,N-dimethylethanamine (DMAZ). In each case, the new technique does a better job than the best current computational methods at accurately matching the experimental data of the HEDM compounds of interest. A case study is provided to help quantify the vehicle level impacts of using HEDM

  4. Combustor design and analysis using the Rocket Combustor Interactive Design (ROCCID) methodology

    NASA Technical Reports Server (NTRS)

    Klem, Mark D.; Pieper, Jerry L.; Walker, Richard E.

    1990-01-01

    The ROCket Combustor Interactive Design (ROCCID) Methodology is a newly developed, interactive computer code for the design and analysis of a liquid propellant rocket combustion chamber. The application of ROCCID to design a liquid rocket combustion chamber is illustrated. Designs for a 50,000 lbf thrust and 1250 psi chamber pressure combustor using liquid oxygen (LOX)RP-1 propellants are developed and evaluated. Tradeoffs between key design parameters affecting combustor performance and stability are examined. Predicted performance and combustion stability margin for these designs are provided as a function of the combustor operating mixture ratio and chamber pressure.

  5. Combustor design and analysis using the ROCket Combustor Interactive Design (ROCCID) Methodology

    NASA Technical Reports Server (NTRS)

    Klem, Mark D.; Pieper, Jerry L.; Walker, Richard E.

    1990-01-01

    The ROCket Combustor Interactive Design (ROCCID) Methodology is a newly developed, interactive computer code for the design and analysis of a liquid propellant rocket combustion chamber. The application of ROCCID to design a liquid rocket combustion chamber is illustrated. Designs for a 50,000 lbf thrust and 1250 psi chamber pressure combustor using liquid oxygen (LOX)RP-1 propellants are developed and evaluated. Tradeoffs between key design parameters affecting combustor performance and stability are examined. Predicted performance and combustion stability margin for these designs are provided as a function of the combustor operating mixture ratio and chamber pressure.

  6. Liquid rocket engine fluid-cooled combustion chambers

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A monograph on the design and development of fluid cooled combustion chambers for liquid propellant rocket engines is presented. The subjects discussed are (1) regenerative cooling, (2) transpiration cooling, (3) film cooling, (4) structural analysis, (5) chamber reinforcement, and (6) operational problems.

  7. Rocket Scientist for a Day: Investigating Alternatives for Chemical Propulsion

    ERIC Educational Resources Information Center

    Angelin, Marcus; Rahm, Martin; Gabrielsson, Erik; Gumaelius, Lena

    2012-01-01

    This laboratory experiment introduces rocket science from a chemistry perspective. The focus is set on chemical propulsion, including its environmental impact and future development. By combining lecture-based teaching with practical, theoretical, and computational exercises, the students get to evaluate different propellant alternatives. To…

  8. Studies of solid propellant combustion with pulsed radiography

    NASA Technical Reports Server (NTRS)

    Godai, T.; Tanemura, T.; Fujiwara, T.; Shimizu, M.

    1987-01-01

    Pulsed radiography was applied to observe solid propellant surface regression during rocket motor operation. Using a 150 KV flash X-ray system manufactured by the Field Emission Corporation and two kinds of film suppliers, images of the propellant surface of a 5 cm diameter end burning rocket motor were recorded on film. The repetition frame rate of 8 pulses per second and the pulse train length of 10 pulses are limited by the capability of the power supply and the heat build up within the X-ray tube, respectively. The experiment demonstrated the effectiveness of pulsed radiography for observing solid propellant surface regression. Measuring the position of burning surface images on film with a microdensitometer, quasi-instantaneous burning rate as a function of pressure and the variation of characteristic velocity with pressure and gas stay time were obtained. Other research items to which pulsed radiography can be applied are also suggested.

  9. Passive propellant system

    NASA Technical Reports Server (NTRS)

    Hess, D. A.; Regnier, W. W.; Jacobs, V. L. (Inventor)

    1979-01-01

    A passive propellant acquisition and feed system is disclosed which acquires and feeds gas-free propellant in low or zero-g environments during orbital maneuvers and retains this propellant under high axially directed acceleration such as may be experienced during launch of a space vehicle and orbit-to-orbit transfer is described. The propellant system includes a dual compartment propellant tank with independent surface tension acquisition channels in each compartment to provide gas-free flow of pressurized liquid propellant from one compartment to the other in one direction only.

  10. 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.

  11. Tandem Air Propellers - II

    NASA Technical Reports Server (NTRS)

    Lesley, E. P.

    1939-01-01

    Tests of three-blade, adjustable-pitch counterrotating tandem model propellers, adjusted to absorb equal power at maximum efficiency of the combination, were made at Stanford University. The aerodynamic characteristics, for blade-angle settings of 15, 25, 35, 45, 55, and 65 degrees at 0.75R of the forward propeller and for diameters spacings of 8-1/2, 15 and 30% were compared with those of three-blade and six-blade propellers of the same blade form. It was found that, in order to realize the condition of equal power at maximum efficiency, the blade angles for the rear propeller must be generally less than for the forward propeller, the difference increasing the blade angle. The tests showed that, at maximum efficiency, the tandem propellers absorb about double the power of three-blade propellers and about 8% more power than six-blade propellers having the pitch of the forward propeller of the tandem combination. The maximum efficiency of the tandem propellers was found to be from 2-15% greater than for six-blade propellers, the difference varying directly with blade angle. It was also found that the maximum efficiency of the tandem propellers was greater than that of a three-blade propeller for blade angles at 0.75R of 25 degrees or more. The difference in maximum efficiency again varied directly with blade angle, being about 9% for 65 degrees at 0.75R.

  12. Rocket effluent: Its ice nucleation activity and related properties

    NASA Technical Reports Server (NTRS)

    Parungo, F. P.; Allee, P. A.

    1978-01-01

    To investigate the possibility of inadvertent weather modification from rocket effluent, aerosol samples were collected from an instrumented aircraft subsequent to the Voyager 1 and 2 launches. The aerosol's morphology, concentration, and size distribution were examined with an electron microscope. The elemental compositions of individual particles were analyzed with an X-ray energy spectrometer. Ice nucleus concentration was measured with a thermal diffusion chamber. The particles' physical and chemical properties were related to their ice nucleation activity. A laboratory experiment on rocket propellant exhaust was conducted under controlled conditions. Both laboratory and field experimental results indicated that rocket propellant exhaust can produce active ice nuclei and modify local weather in suitable meteorological conditions.

  13. Rocket effluent - Its ice nucleation activity and related properties

    NASA Technical Reports Server (NTRS)

    Parungo, F. P.; Allee, P. A.

    1978-01-01

    To investigate the possibility of inadvertent weather modification from rocket effluent, aerosol samples were collected from an instrumented aircraft subsequent to the Voyager I and II launches. The aerosol's morphology, concentration and size distribution were examined with an electron microscope. The elemental compositions of individual particles were analyzed with an X-ray energy spectrometer. Ice nucleus concentration was measured with a subfreezing thermal diffusion chamber. The particles' physical and chemical properties were related to their ice nucleation activity. A laboratory experiment on rocket propellant exhaust was conducted under controlled conditions. Both laboratory and field experimental results indicated that rocket propellant exhaust can produce active ice nuclei. Their consequences for potential inadvertant weather modification demand additional study.

  14. Characterizing high-energy-density propellants for space propulsion applications

    NASA Astrophysics Data System (ADS)

    Kokan, Timothy S.; Olds, John R.; Seitzman, Jerry M.; Ludovice, Peter J.

    2009-10-01

    A technique for computationally determining the thermophysical properties of high-energy-density matter (HEDM) propellants is presented. HEDM compounds are of interest in the liquid rocket engine industry due to their high density and high energy content relative to existing industry-standard propellants. In order to accurately model rocket engine performance, cost and weight in a conceptual design environment, several thermodynamic and physical properties are required over a range of temperatures and pressures. The approach presented here combines quantum mechanical and molecular dynamic (MD) calculations and group additivity methods. A method for improving the force field model coefficients used in the MD is included. This approach is used to determine thermophysical properties for two HEDM compounds of interest: quadricyclane and 2-azido-N,N-dimethylethanamine (DMAZ). The modified force field approach provides results that more accurately match experimental data than the unmodified approach. Launch vehicle and Lunar lander case studies are presented to quantify the system level impact of employing quadricyclane and DMAZ rather than industry standard propellants. In both cases, the use of HEDM propellants provides reductions in vehicle mass compared to industry standard propellants. The results demonstrate that HEDM propellants can be an attractive technology for future launch vehicle and Lunar lander applications.

  15. 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.

  16. 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.

  17. Dynamic analysis of an open-cycle gas-core nuclear rocket

    NASA Technical Reports Server (NTRS)

    Turney, G. E.; Kieffer, A. W.

    1973-01-01

    Reactivity and flow disturbances were used to investigate the transient response of a conceptual open cycle gas core nuclear rocket engine. The disturbances were made with the system initially operating at its steady state design point. Results of the study show that the feedbacks associated with the propellant density and propellant temperature have a dominant effect on the response of the system. Furthermore, there appears to be a rather limited range of values of these propellant feedback coefficients for which the gas core nuclear rocket has a stable response. The system was rather insensitive to a fuel flow rate disturbance, whereas a similar disturbance in the propellant flow rate caused large changes in reactor power. For a similar disturbance in the propellant flow rate caused large changes in reactor power. For most reactivity and flow rate disturbances, the response showed oscillations of various intensity.

  18. 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.

  19. 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.

  20. 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.

  1. Contained rocket motor burn demonstrations in X-tunnel: Final report for the DoD/DOE Joint Demilitarization Technology Program

    SciTech Connect

    S. W. Allendorf; B. W. Bellow; R. f. Boehm

    2000-05-01

    Three low-pressure rocket motor propellant burn tests were performed in a large, sealed test chamber located at the X-tunnel complex on the Department of Energy's Nevada Test Site in the period May--June 1997. NIKE rocket motors containing double base propellant were used in two tests (two and four motors, respectively), and the third test used two improved HAWK rocket motors containing composite propellant. The preliminary containment safety calculations, the crack and burn procedures used in each test, and the results of various measurements made during and after each test are all summarized and collected in this document.

  2. Passive propellant system

    NASA Technical Reports Server (NTRS)

    Hess, D. A.; Regnier, W. W.; Jacobs, V. L. (Inventor)

    1978-01-01

    The system utilizes a spherical tank structure A separated into two equal volume compartments by a flat bulkhead B. Each compartment has four similar gallery channel legs located in the principal vehicle axes, ensuring that bulk propellant will contact at least one gallery leg during vehicle maneuvers. The forward compartment gallery channel legs collect propellant and feed it into the aft compartment through communication screens which protrude into the aft compartment. The propellant is then collected by the screened gallery channels in the aft compartment and supplied to the propellant outlet. The invention resides in the independent gallery assembly and screen structure by means of which propellant flow from forward to aft compartments is maintained. Liquid surface tension of the liquid on the screens is used to control liquid flow. The system provides gas-free propellants in low or zero-g environments regardless of axial accelerations and propellant orientation in bulk regions of the vessel.

  3. Metallized Gelled Propellant Heat Transfer Tests Analyzed

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan A.

    1997-01-01

    A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted at the NASA Lewis Research Center. These experiments used a small 20- to 40-lbf thrust engine composed of a modular injector, an igniter, a chamber, and a nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt % loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each channel used water flow to carry heat away from the chamber and the attached thermocouples; flow meters allowed heat flux estimates at each of the 31 stations.

  4. Propellant Feed Subsystem for the X-34 Main Propulsion System

    NASA Technical Reports Server (NTRS)

    McDonald, J. P.; Minor, R. B.; Knight, K. C.; Champion, R. H., Jr.; Russell, F. J., Jr.

    1998-01-01

    The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major NMS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. This paper reviews major design and analysis aspects of the X-34 propellant feed subsystem of the X-34 Main Propulsion System. Topics include system requirements, system design, the integration of flight and feed system performance, propellant acquisition at engine start, and propellant tank terminal drain.

  5. Modal survey of the space shuttle solid rocket motor using multiple input methods

    NASA Technical Reports Server (NTRS)

    Brillhart, Ralph; Hunt, David L.; Jensen, Brent M.; Mason, Donald R.

    1987-01-01

    The ability to accurately characterize propellant in a finite element model is a concern of engineers tasked with studying the dynamic response of the Space Shuttle Solid Rocket Motor (SRM). THe uncertainties arising from propellant characterization through specimem testing led to the decision to perform a model survey and model correlation of a single segment of the Shuttle SRM. Multiple input methods were used to excite and define case/propellant modes of both an inert segment and, later, a live propellant segment. These tests were successful at defining highly damped, flexible modes, several pairs of which occured with frequency spacing of less than two percent.

  6. 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.

  7. 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.

  8. Inflatable device for excluding sea water from a rocket motor

    NASA Astrophysics Data System (ADS)

    Purser, Michael J.

    1992-06-01

    An inflatable water exclusion device for protecting the rocket motor of a solid-propellant-powered missile, launched from below the surface of the sea by means of launching gas, from being damaged by a spout of water spurting up into the atmosphere. The device consists of a round base plate that attaches to an expellable baffle assembly located within the throat of the rocket motor nozzle, and an inflatable elastomer-coated cloth bag that is attached to the rim of the base plate. As the missile ascends toward the surface of the sea, launching gas that was originally contained within the voids of the rocket motor expands and streams out through the nozzle thereby causing the bag to inflate. After the missile has breached the surface of the sea the rocket motor ignites, thereby causing the baffle assembly and the inflatable water exclusion device attached thereto to be expelled from the nozzle and fall back into the sea.

  9. Airbreathing/Rocket Single-Stage-to-Orbit Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.

    1995-01-01

    A definitive design/performance study was performed on a single-stage-to-orbit (SSTO) airbreathing propelled orbital vehicle with rocket propulsion augmentation in the Access to Space activities during 1993. A credible reference design was established, but by no means an optimum. The results supported the viability of SSTO airbreathing/rocket vehicles for operational scenarios and indicated compelling reasons to continue to explore the design matrix. This paper will (1) summarize the Access to Space design activity from the SSTO airbreathing/rocket perspective, (2) present an airbreathing/rocket SSTO design matrix established for continued optimization of the design space, and (3) focus on the compelling reasons for airbreathing vehicles in Access to Space scenarios.

  10. 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.

  11. The effects of solid rocket motor effluents on selected surfaces and solid particle size, distribution, and composition for simulated shuttle booster separation motors

    NASA Technical Reports Server (NTRS)

    Jex, D. W.; Linton, R. C.; Russell, W. M.; Trenkle, J. J.; Wilkes, D. R.

    1976-01-01

    A series of three tests was conducted using solid rocket propellants to determine the effects a solid rocket plume would have on thermal protective surfaces (TPS). The surfaces tested were those which are baselined for the shuttle vehicle. The propellants used were to simulate the separation solid rocket motors (SSRM) that separate the solid rocket boosters (SRB) from the shuttle launch vehicle. Data cover: (1) the optical effects of the plume environment on spacecraft related surfaces, and (2) the solid particle size, distribution, and composition at TPS sample locations.

  12. Fluid-solid coupled simulation of the ignition transient of solid rocket motor

    NASA Astrophysics Data System (ADS)

    Li, Qiang; Liu, Peijin; He, Guoqiang

    2015-05-01

    The first period of the solid rocket motor operation is the ignition transient, which involves complex processes and, according to chronological sequence, can be divided into several stages, namely, igniter jet injection, propellant heating and ignition, flame spreading, chamber pressurization and solid propellant deformation. The ignition transient should be comprehensively analyzed because it significantly influences the overall performance of the solid rocket motor. A numerical approach is presented in this paper for simulating the fluid-solid interaction problems in the ignition transient of the solid rocket motor. In the proposed procedure, the time-dependent numerical solutions of the governing equations of internal compressible fluid flow are loosely coupled with those of the geometrical nonlinearity problems to determine the propellant mechanical response and deformation. The well-known Zeldovich-Novozhilov model was employed to model propellant ignition and combustion. The fluid-solid coupling interface data interpolation scheme and coupling instance for different computational agents were also reported. Finally, numerical validation was performed, and the proposed approach was applied to the ignition transient of one laboratory-scale solid rocket motor. For the application, the internal ballistics were obtained from the ground hot firing test, and comparisons were made. Results show that the integrated framework allows us to perform coupled simulations of the propellant ignition, strong unsteady internal fluid flow, and propellant mechanical response in SRMs with satisfactory stability and efficiency and presents a reliable and accurate solution to complex multi-physics problems.

  13. Minuteman 3: Stage 3 propellant fire characterization

    SciTech Connect

    Diaz, J C

    1994-06-20

    We have completed an experimental program to diagnose and characterize the thermal environment of a solid rocket propellant fire burning in ambient atmospheric conditions. This work has been conducted as part of the Defense Nuclear Agency`s probabilistic risk assessment (PRA) of the Minuteman III (MMIII) weapon system. The goals of this study are two-fold; (1) to provide a description of a propellant fire in sufficient detail so as to allow system response models to predict the outcome of various hypothetical accident, scenarios and, (2) to identify diagnostics that could be used in a large-scale test fire of MMIII stage 3 motor. The study has been performed burning SRAM II and MMIII stage 3 propellant (ANB-3066), in chimneys ranging in size from 18 cm to 53 cm (twenty-one inches) in diameter. Several methods have been used to determine and confirm temperature measurements including thermometry, radiometry, and ultrasonic thermal sensing. Temperature profiles with peaks in excess of 2300{degree} C have been measured. Temperature measurements have been used in conjunction with inverse modeling to determine heat flux characteristics. The regression rate for ANB-3066 (under ambient conditions) has also been determined. Finally, at a very cursory level, we have studied the characteristics of aluminum oxide deposits as well as some materials responses to these fires. We have also addressed the initial efforts on development of diagnostics, problems encountered in controlling the burning of propellants, in taking radiometric measurements, and the survivability of materials in the fire.

  14. Demonstration of a sterilizable solid rocket motor system

    NASA Technical Reports Server (NTRS)

    Mastrolia, E. J.; Santerre, G. M.; Lambert, W. L.

    1975-01-01

    A solid propellant rocket motor containing 60.9 Kg (134-lb) of propellant was successfully static fired after being subjected to eight heat sterilization cycles (three 54-hour cycles plus five 40-hour cycles) at 125 C (257 F). The test motor, a modified SVM-3 chamber, incorporated a flexible grain retention system of EPR rubber to relieve thermal shrinkage stresses. The propellant used in the motor was ANB-3438, and 84 wt% solids system (18 wt% aluminum) containing 66 wt% stabilized ammonium perchlorate oxidizer and a saturated hydroxylterminated polybutadiene binder. Bonding of the propellant to the EPR insulation (GenGard V-4030) was provided by the use of SD-886, an epoxy urethane restriction.

  15. 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.

  16. 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.

  17. An Analysis of Rocket Propulsion Testing Costs

    NASA Technical Reports Server (NTRS)

    Ramirez, Carmen; Rahman, Shamim

    2010-01-01

    The primary mission at NASA Stennis Space Center (SSC) is rocket propulsion testing. Such testing is commonly characterized as one of two types: production testing for certification and acceptance of engine hardware, and developmental testing for prototype evaluation or research and development (R&D) purposes. For programmatic reasons there is a continuing need to assess and evaluate the test costs for the various types of test campaigns that involve liquid rocket propellant test articles. Presently, in fact, there is a critical need to provide guidance on what represents a best value for testing and provide some key economic insights for decision-makers within NASA and the test customers outside the Agency. Hence, selected rocket propulsion test databases and references have been evaluated and analyzed with the intent to discover correlations of technical information and test costs that could help produce more reliable and accurate cost projections in the future. The process of searching, collecting, and validating propulsion test cost information presented some unique obstacles which then led to a set of recommendations for improvement in order to facilitate future cost information gathering and analysis. In summary, this historical account and evaluation of rocket propulsion test cost information will enhance understanding of the various kinds of project cost information; identify certain trends of interest to the aerospace testing community.

  18. 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.

  19. Rocket Engine Innovations Advance Clean Energy

    NASA Technical Reports Server (NTRS)

    2012-01-01

    During launch countdown, at approximately T-7 seconds, the Space Shuttle Main Engines (SSMEs) roar to life. When the controllers indicate normal operation, the solid rocket boosters ignite and the shuttle blasts off. Initially, the SSMEs throttle down to reduce stress during the period of maximum dynamic pressure, but soon after, they throttle up to propel the orbiter to 17,500 miles per hour. In just under 9 minutes, the three SSMEs burn over 1.6 million pounds of propellant, and temperatures inside the main combustion chamber reach 6,000 F. To cool the engines, liquid hydrogen circulates through miles of tubing at -423 F. From 1981to 2011, the Space Shuttle fleet carried crew and cargo into orbit to perform a myriad of unprecedented tasks. After 30 years and 135 missions, the feat of engineering known as the SSME boasted a 100-percent flight success rate.

  20. Recent Advances and Applications in Cryogenic Propellant Densification Technology

    NASA Technical Reports Server (NTRS)

    Tomsik, Thomas M.

    2000-01-01

    This purpose of this paper is to review several historical cryogenic test programs that were conducted at the NASA Glenn Research Center (GRC), Cleveland, Ohio over the past fifty years. More recently these technology programs were intended to study new and improved denser forms of liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic rocket fuels. Of particular interest are subcooled cryogenic propellants. This is due to the fact that they have a significantly higher density (eg. triple-point hydrogen, slush etc.), a lower vapor pressure and improved cooling capacity over the normal boiling point cryogen. This paper, which is intended to be a historical technology overview, will trace the past and recent development and testing of small and large-scale propellant densification production systems. Densifier units in the current GRC fuels program, were designed and are capable of processing subcooled LH2 and L02 propellant at the X33 Reusable Launch Vehicle (RLV) scale. One final objective of this technical briefing is to discuss some of the potential benefits and application which propellant densification technology may offer the industrial cryogenics production and end-user community. Density enhancements to cryogenic propellants (LH2, LO2, CH4) in rocket propulsion and aerospace application have provided the opportunity to either increase performance of existing launch vehicles or to reduce the overall size, mass and cost of a new vehicle system.

  1. 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.

  2. Combustion dynamics in liquid rocket engines

    NASA Technical Reports Server (NTRS)

    Mclain, W. H.

    1971-01-01

    A chemical analysis of the emission and absorption spectra in the combustion chamber of a nitrogen tetroxide/aerozine-50 rocket engine was conducted. Measurements were made under conditions of preignition, ignition, and post combustion operating periods. The cause of severe ignition overpressures sporadically observed during the vacuum startup of the Apollo reaction control system engine was investigated. The extent to which residual propellants or condensed intermediate reaction products remain after the engine has been operated in a pulse mode duty cycle was determined.

  3. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Mellish, J. A.

    1981-01-01

    Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

  4. 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.

  5. Propeller tip vortex interactions

    NASA Technical Reports Server (NTRS)

    Johnston, Robert T.; Sullivan, John P.

    1990-01-01

    Propeller wakes interacting with aircraft aerodynamic surfaces are a source of noise and vibration. For this reason, flow visualization work on the motion of the helical tip vortex over a wing and through the second stage of a counterrotation propeller (CRP) has been pursued. Initially, work was done on the motion of a propeller helix as it passes over the center of a 9.0 aspect ratio wing. The propeller tip vortex experiences significant spanwise displacements when passing across a lifting wing. A stationary propeller blade or stator was installed behind the rotating propeller to model the blade vortex interaction in a CRP. The resulting vortex interaction was found to depend on the relative vortex strengths and vortex sign.

  6. Propeller design by optimization

    NASA Technical Reports Server (NTRS)

    Rizk, M. H.; Jou, W.-H.

    1986-01-01

    The feasibility of designing propellers by an optimization procedure is investigated. A scheme, which solves the full potential flow equation about a propeller by line relaxation, is modified so that the iterative solutions of the flow equation and the design parameters are updated simultaneously. Some technical problems in using optimization for designing propellers with maximum efficiency are identified. Approaches for overcoming these problems are presented.

  7. Analysis of quasi-hybrid solid rocket booster concepts for advanced earth-to-orbit vehicles

    NASA Technical Reports Server (NTRS)

    Zurawski, Robert L.; Rapp, Douglas C.

    1987-01-01

    A study was conducted to assess the feasibility of quasi-hybrid solid rocket boosters for advanced Earth-to-orbit vehicles. Thermochemical calculations were conducted to determine the effect of liquid hydrogen addition, solids composition change plus liquid hydrogen addition, and the addition of an aluminum/liquid hydrogen slurry on the theoretical performance of a PBAN solid propellant rocket. The space shuttle solid rocket booster was used as a reference point. All three quasi-hybrid systems theoretically offer higher specific impulse when compared with the space shuttle solid rocket boosters. However, based on operational and safety considerations, the quasi-hybrid rocket is not a practical choice for near-term Earth-to-orbit booster applications. Safety and technology issues pertinent to quasi-hybrid rocket systems are discussed.

  8. Cryogenic Propellant Densification Study

    NASA Technical Reports Server (NTRS)

    Ewart, R. O.; Dergance, R. H.

    1978-01-01

    Ground and vehicle system requirements are evaluated for the use of densified cryogenic propellants in advanced space transportation systems. Propellants studied were slush and triple point liquid hydrogen, triple point liquid oxygen, and slush and triple point liquid methane. Areas of study included propellant production, storage, transfer, vehicle loading and system requirements definition. A savings of approximately 8.2 x 100,000 Kg can be achieved in single stage to orbit gross liftoff weight for a payload of 29,484 Kg by utilizing densified cryogens in place of normal boiling point propellants.

  9. Propeller flow visualization techniques

    NASA Technical Reports Server (NTRS)

    Stefko, G. L.; Paulovich, F. J.; Greissing, J. P.; Walker, E. D.

    1982-01-01

    Propeller flow visualization techniques were tested. The actual operating blade shape as it determines the actual propeller performance and noise was established. The ability to photographically determine the advanced propeller blade tip deflections, local flow field conditions, and gain insight into aeroelastic instability is demonstrated. The analytical prediction methods which are being developed can be compared with experimental data. These comparisons contribute to the verification of these improved methods and give improved capability for designing future advanced propellers with enhanced performance and noise characteristics.

  10. A theoretical evaluation of aluminum gel propellant two-phase flow losses on vehicle performance

    NASA Technical Reports Server (NTRS)

    Mueller, Donn C.; Turns, Stephen R.

    1993-01-01

    A one-dimensional model of a hydrocarbon/Al/O2(gaseous) fueled rocket combustion chamber was developed to study secondary atomization effects on propellant combustion. This chamber model was coupled with a two dimensional, two-phase flow nozzle code to estimate the two-phase flow losses associated with solid combustion products. Results indicate that moderate secondary atomization significantly reduces propellant burnout distance and Al2O3 particle size; however, secondary atomization provides only moderate decreases in two-phase flow induced I(sub sp) losses. Despite these two-phase flow losses, a simple mission study indicates that aluminum gel propellants may permit a greater maximum payload than the hydrocarbon/O2 bi-propellant combination for a vehicle of fixed propellant volume. Secondary atomization was also found to reduce radiation losses from the solid combustion products to the chamber walls, primarily through reductions in propellant burnout distance.

  11. Concepts for the design of an antimatter annihilation rocket

    NASA Technical Reports Server (NTRS)

    Morgan, D. L., Jr.

    1982-01-01

    Matter-antimatter annihilation is considered for spacecraft propulsion. Annihilation produces considerably more energy per unit mass of propellant than any other known means of energy production. An antimatter annihilation rocket requires several systems and components that are unique to its nature. Among these are an antimatter storage system, a means to extract the antimatter from storage, a system to transport the antimatter to the rocket engine, and the engine wherein annihilation occurs and thrust is produced. Design concepts of these systems and components are presented and discussed.

  12. 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.

  13. Miniature Rocket Motor for Aircraft Stall/Spin Recovery

    NASA Technical Reports Server (NTRS)

    Lucy, M. H.

    1985-01-01

    Design accommodates different thrust levels and burn times with minimum weight. Different thrust levels achieved by substituting other propellants of different diameter and burn-rate characteristics. Different burn times achieved by simply changing length of grain/tube assembly. Grain bond material also acts as insulator for fiberglass tube. Rocket motor attached to aircraft model and ignited from radio-controlled 4.8-volt power source. Device provides more than twice energy available in previous designs at only 60 percent of weight. Rocket motor used to identify energy requirements for aircraft stall/spin recovery positive propulsion system.

  14. 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.

  15. Questions of testing rate and flexibility of rocket test benches, discussed on the basis of the test benches of Nitrochemie GMBH in Aschau

    NASA Technical Reports Server (NTRS)

    LEGRAND

    1987-01-01

    The rocket test benches are used to study burnup behavior by various methods. In the first ten months of 1966, 1578 shots were performed to test propellants, and 920 to test 14 thrust and pressure measurement projects.

  16. Ice nucleus activity measurements of solid rocket motor exhaust particles

    NASA Technical Reports Server (NTRS)

    Keller, V. W. (Compiler)

    1986-01-01

    The ice Nucleus activity of exhaust particles generated from combustion of Space Shuttle propellant in small rocket motors has been measured. The activity at -20 C was substantially lower than that of aerosols generated by unpressurized combustion of propellant samples in previous studies. The activity decays rapidly with time and is decreased further in the presence of moist air. These tests corroborate the low effectivity ice nucleus measurement results obtained in the exhaust ground cloud of the Space Shuttle. Such low ice nucleus activity implies that Space Shuttle induced inadvertent weather modification via an ice phase process is extremely unlikely.

  17. Worldwide Space Launch Vehicles and Their Mainstage Liquid Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim A.

    2010-01-01

    Space launch vehicle begins with a basic propulsion stage, and serves as a missile or small launch vehicle; many are traceable to the 1945 German A-4. Increasing stage size, and increasingly energetic propulsion allows for heavier payloads and greater. Earth to Orbit lift capability. Liquid rocket propulsion began with use of storable (UDMH/N2O4) and evolved to high performing cryogenics (LOX/RP, and LOX/LH). Growth versions of SLV's rely on strap-on propulsive stages of either solid propellants or liquid propellants.

  18. Erosive burning threshold conditions in solid rocket motors

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Cohen, N. S.

    1989-01-01

    Erosive burning has been predicted to be enhanced by factors that increase the level of turbulence close to the propellant surface, such a high cross flow velocity, low surface blowing rate, propellant surface roughness, and adverse pressure gradient. A study is reported which was carried out with the objective of measuring the effects of these parameters on the scaling to larger rocket motor sizes of the transition to, or threshold conditions for, erosive burning rate augmentation. The results are used to develop a scaling criterion for the threshold conditions for erosive burning.

  19. The use of electrical discharge for ignition and control of combustion of solid propellants

    NASA Technical Reports Server (NTRS)

    Tachibana, Takeshi; Kobayashi, Tsuruo; Matsuda, Takashi; Kimura, Itsuro

    1987-01-01

    As the first step of the study of the combustion control of solid propellants by electrical discharges, the effects of an arc discharge, which flows along the burning surface, on the burning rate and on the increase of enthalpy of the combustion product were investigated. For specially devised composite propellants, which are composed of Al and Teflon powders, it was shown that the combination can be controlled by an arc discharge; the combustion continues when the arc discharge is applied and is interrupted when the arc discharge breaks. In the present investigation, it was also shown that an arc discharge coupled with a high-frequency electrical discharge has potential as an effective ignition method for solid propellants. For the application of this type of combustion control to an ignitor for a solid propellant rocket motor or to a control rocket motor, this method lacks flexibility in the configuration scale and needs relatively high electric power at the present stage.

  20. Injection dynamics of gelled propellants

    NASA Astrophysics Data System (ADS)

    Yoon, Changjin

    Gel propellants have been recognized as attractive candidates for future propulsion systems due to the reduced tendency to spill and the energy advantages over solid propellants. One of strong benefits emphasized in gel propellant applications is a throttling capability, but the accurate flow control is more complicated and difficult than with conventional Newtonian propellants because of the unique rheological behaviors of gels. This study is a computational effort directed to enhance understanding of the injector internal flow characteristics for gel propellants under rocket injection conditions. In simulations, the emphasized rheology is a shear-thinning which represents a viscosity decrease with increasing a shear rate. It is described by a generalized Newtonian fluid constitutive equation and Carreau-Yasuda model. Using this rheological model, two injection schemes are considered in the present study: axially-fed and cross-fed injection for single-element and multi-element impinging injectors, respectively. An axisymmetric model is developed to describe the axially-fed injector flows and fully three-dimensional model is utilized to simulate cross-fed injector flows. Under axially-fed injection conditions investigated, three distinct modes, an unsteady, steady, and hydraulic flip mode, are observed and mapped in terms of Reynolds number and orifice design. In an unsteady mode, quasi-periodic oscillations occur near the inlet lip leading mass pulsations and viscosity fluctuations at the orifice exit. This dynamic behavior is characterized using a time-averaged discharge coefficient, oscillation magnitude and frequency by a parametric study with respect to an orifice design, Reynolds number and rheology. As a result, orifice exit flows for gel propellants appear to be significantly influenced by a viscous damping and flow resistance due to a shear thinning behavior and these are observed in each factors considered. Under conditions driven by a manifold crossflow

  1. Pellet bed reactor for nuclear propelled vehicles: Part 1: Reactor technology

    NASA Technical Reports Server (NTRS)

    El-Genk, Mohamed S.

    1991-01-01

    The pellet bed reactor (PBR) for nuclear propelled vehicles is briefly discussed. Much of the information is given in viewgraph form. Viewgraphs include information on the layout for a Mars mission using a PBR nuclear thermal rocket, the rocket reactor layout, the fuel pellet design, materials compatibility, fuel microspheres, microsphere coating, melting points in quasibinary systems, stress analysis of microspheres, safety features, and advantages of the PBR concept.

  2. Manned Mars missions using propellant from space

    SciTech Connect

    Zuppero, A.C.; Olson, T.S. ); Redd, L.R. )

    1993-01-10

    .A recent discovery (8/14/92) of a near-earth object containing materials potentially useful for space activities could perhaps change the entire way humans access and operate in space. A near-Earth object ([number sign]4015, 1979 VA, comet Wilson-Harrington) contains water ice that could be used for space propulsion. In addition, this type of object may contain structural and lifesustaining materials (complex hydrocarbons, ammonia and/or bound nitrogen compounds) for space structures, manned planetary bases, or planetary surface terraforming. The retrieval and utilization of rocket propellant from near-Earth objects, for manned Mars missions in particular, has been investigated and the benefits of this scenario to over performing a Mars mission with terrestrial propellants have been documented. The results show water extracted from these objects and retrieved to Earth orbit for use in going to Mars may actually enable manned Mars exploration by reducing the number of Heavy Lift Launch Vehicle (HLLV) flights or eliminating the need for HLLV's altogether. The mission can perhaps be supported with existing launch vehicles and not required heavy lift capability. Also, the development of a nuclear thermal rocket for this alternate approach may be simplified substantially by reducing the operating temperature required.

  3. Solid Hydrogen Formed for Atomic Propellants

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan A.

    2000-01-01

    Several experiments on the formation of solid hydrogen particles in liquid helium were recently conducted at the NASA Glenn Research Center at Lewis Field. The solid hydrogen experiments are the first step toward seeing these particles and determining their shape and size. The particles will ultimately store atoms of boron, carbon, or hydrogen, forming an atomic propellant. Atomic propellants will allow rocket vehicles to carry payloads many times heavier than possible with existing rockets or allow them to be much smaller and lighter. Solid hydrogen particles are preferred for storing atoms. Hydrogen is generally an excellent fuel with a low molecular weight. Very low temperature hydrogen particles (T < 4 K) can prevent the atoms from recombining, making it possible for their lifetime to be controlled. Also, particles that are less than 1 mm in diameter are preferred because they can flow easily into a pipe when suspended in liquid helium. The particles and atoms must remain at this low temperature until the fuel is introduced into the engine combustion (or recombination) chamber. Experiments were, therefore, planned to look at the particles and observe their formation and any changes while in liquid helium.

  4. Ignition transient analysis of a solid rocket motor using a one dimensional two fluid model

    NASA Astrophysics Data System (ADS)

    Pardue, Byron A.; Han, Samuel S.

    1992-07-01

    A one dimensional two fluid numerical model has been used to study the ignition transient stage of a Space Shuttle solid rocket motor. During the ignition phase of a solid rocket motor a pressure transient is induced by complex transport processes involving the igniter gas heat transfer to the propellant, chemical reactions at the propellant surface, and the interaction of the fluid with the attached rocket nozzle. One dimensional models used in the past neglected the aluminum oxide particles which are present in the combustion gases. The current model uses the IPSA (Inter-Phase-Slip-Algorithm) to solve the transient compressible flow equations for the rocket chamber and attached nozzle. Numerical results for head end pressure changes and overall thrust are compared with both measurement data and predictions of a one dimensional one fluid model.

  5. Accelerated Aging of Lead-Free Propellant

    NASA Technical Reports Server (NTRS)

    Furrow, Keith W.; Jervey, David D.

    2000-01-01

    Following higher than expected 2-NDPA depletion rates in a lead-free doublebase formulation (RPD-422), an accelerated aging study was conducted to verify the depletion rates. A test plan was prepared to compare the aging characteristics of lead-free propellant and NOSIH-AA2. The study was also designed to determine which lead-free ballistic modifiers accelerated 2-NDPA depletion. The increased depletion rate occurred in propellants containing monobasic copper salicylate. Four lead-free propellants were then formulated to improved aging characteristics over previous lead-free propellant formulations. The new formulations reduced or replaced the monobasic copper salicylate. The new formulations had improved aging characteristics. Their burn rates, however, were unacceptable for use in a 2.75 inch rocket. To compare aging characteristics, stabilizer depletion rates of RPD-422, AA2, M28, and RLC 470/6A were measured or taken from the literature. The data were fit to a kinetic model. The model contained first and zero order terms which allowed the stabilizer concentration to go to zero. In the model, only the concentration of the primary stabilizer was considered. Derivatives beyond the first nitrated or nitroso derivative of 2-NPDA were not considered. The rate constants were fit to the Arrhenius equation and extrapolated to lower temperatures. The time to complete stabilizer depletion was estimated using the kinetic model. The four propellants were compared and the RPD-422 depleted faster at 45 C than both A22 and M28. These types of predictions depend on the validity of the model and on confidence in the Arrhenius relationship holding at lower temperatures. At 45 C, the zero order portion of the model dominates the depletion rate.

  6. The Guggenheim Aeronautics Laboratory at Caltech and the creation of the modern rocket motor (1936-1946): How the dynamics of rocket theory became reality

    NASA Astrophysics Data System (ADS)

    Zibit, Benjamin Seth

    This thesis explores and unfolds the story of discovery in rocketry at The California Institute of Technology---specifically at Caltech's Guggenheim Aeronautics Laboratory---in the 1930s and 1940s. Caltech was home to a small group of engineering students and experimenters who, beginning in the winter of 1935--1936, formed a study and research team destined to change the face of rocket science in the United States. The group, known as the Guggenheim Aeronautics Laboratory (GALCIT, for short) Rocket Research Group, invented a new type of solid-rocket propellant, made distinct and influential discoveries in the theory of rocket combustion and design, founded the Jet Propulsion Laboratory, and incorporated the first American industrial concern devoted entirely to rocket motor production: The Aerojet Corporation. The theoretical work of team members, Frank Malina, Hsueh-shen Tsien, Homer J. Stewart, and Mark Mills, is examined in this thesis in detail. The author scrutinizes Frank Malina's doctoral thesis (both its assumptions and its mathematics), and finds that, although Malina's key assertions, his formulae, hold, his work is shown to make key assumptions about rocket dynamics which only stand the test of validity if certain approximations, rather than exact measurements, are accepted. Malina studied the important connection between motor-nozzle design and thrust; in his Ph.D. thesis, he developed mathematical statements which more precisely defined the design/thrust relation. One of Malina's colleagues on the Rocket Research Team, John Whiteside Parsons, created a new type of solid propellant in the winter of 1941--1942. This propellant, known as a composite propellant (because it simply was a relatively inert amalgam of propellant and oxidizer in non-powder form), became the forerunner of all modern solid propellants, and has become one of the seminal discoveries in the field of Twentieth Century rocketry. The latter chapters of this dissertation discuss the

  7. An Improved Model of Cryogenic Propellant Stratification in a Rotating, Reduced Gravity Environment

    NASA Technical Reports Server (NTRS)

    Oliveira, Justin; Kirk, Daniel R.; Schallhorn, Paul A.; Piquero, Jorge L.; Campbell, Mike; Chase, Sukhdeep

    2007-01-01

    This paper builds on a series of analytical literature models used to predict thermal stratification within rocket propellant tanks. The primary contribution to the literature is to add the effect of tank rotation and to demonstrate the influence of rotation on stratification times and temperatures. This work also looks levels of thermal stratification for generic propellant tanks (cylindrical shapes) over a parametric range of upper-stage coast times, heating levels, rotation rates, and gravity levels.

  8. Current status of free radicals and electronically excited metastable species as high energy propellants

    NASA Technical Reports Server (NTRS)

    Rosen, G.

    1973-01-01

    A survey is presented of free radicals and electronically excited metastable species as high energy propellants for rocket engines. Nascent or atomic forms of diatomic gases are considered free radicals as well as the highly reactive diatomic triatomic molecules that posess unpaired electrons. Manufacturing and storage problems are described, and a review of current experimental work related to the manufacture of atomic hydrogen propellants is presented.

  9. A study of performance and cost improvement potential of the 120 inch (3.05 m) diameter solid rocket motor. Volume 1: Summary report

    NASA Technical Reports Server (NTRS)

    Backlund, S. J.; Rossen, J. N.

    1971-01-01

    A parametric study of ballistic modifications to the 120 inch diameter solid propellant rocket engine which forms part of the Air Force Titan 3 system is presented. 576 separate designs were defined and 24 were selected for detailed analysis. Detailed design descriptions, ballistic performance, and mass property data were prepared for each design. It was determined that a relatively simple change in design parameters could provide a wide range of solid propellant rocket engine ballistic characteristics for future launch vehicle applications.

  10. 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.

  11. Mobile propeller dynamometer validation

    NASA Astrophysics Data System (ADS)

    Morris, Mason Wade

    With growing interest in UAVs and OSU's interest in propeller performance and manufacturing, evaluating UAV propeller and propulsion system performance has become essential. In attempts to evaluate these propellers a mobile propeller dynamometer has been designed, built, and tested. The mobile dyno has been designed to be cost effective through the ability to load it into the back of a test vehicle to create simulated forward flight characteristics. This allows much larger propellers to be dynamically tested without the use of large and expensive wind tunnels. While evaluating the accuracy of the dyno, several improvements had to be made to get accurate results. The decisions made to design and improve the mobile propeller dyno will be discussed along with attempts to validate the dyno by comparing its results against known sources. Another large part of assuring the accuracy of the mobile dyno is determining if the test vehicle will influence the flow going into the propellers being tested. The flow into the propeller needs to be as smooth and uniform as possible. This is determined by characterizing the boundary layer and accelerated flow over the vehicle. This evaluation was accomplished with extensive vehicle aerodynamic measurements with the use of full-scale tests using a pitot-rake and the actual test vehicle. Additional tests were conducted in Oklahoma State University's low speed wind tunnel with a 1/8-scale model using qualitative flow visualization with smoke. Continuing research on the mobile dyno will be discussed, along with other potential uses for the dyno.

  12. Liquid propellant densification

    NASA Technical Reports Server (NTRS)

    Lak, Tibor I. (Inventor); Petrilla, Steve P. (Inventor); Lozano, Martin E. (Inventor)

    1997-01-01

    Super cooling the cryogenic liquid propellant in a vehicle propellant tank densities the propellant allowing the vehicle propellant tank to carry more fuel in the same volume tank while lowering the vapor pressure and thus the tank operating pressure. Lowering the tank operating pressure reduces the stress and therefore allows the walls of the tank to be thinner. Both the smaller tank volume and thinner tank wall results in an overall smaller and lighter vehicle with increased payload capability. The cryogenic propellant can be supercooled well below the normal boiling point temperature level by transporting the liquid propellant from the vehicle tanks to a ground based cooling unit which utilizes a combination of heat exchanger and compressor. The compressor lowers the coolant fluid bath pressure resulting in a low temperature boiling liquid which is subsequently used to cool the recirculating liquid. The cooled propellant is then returned to the vehicle propellant tank. In addition to reducing the vehicle size and weight the invention also allows location of the vent valve on the ground, elimination of on-board recirculation pumps or bleed systems, smaller and lighter engine pumps and valves, lighter and more stable ullage gas, and significant reduction in tank fill operation. All of these mentioned attributes provide lower vehicle weight and cost.

  13. Return of the propeller

    SciTech Connect

    Not Available

    1987-05-01

    Resurrecting the propeller-driven airplane could help save fuel if there is another oil crisis like in the 1970s. This article discusses the new propeller engine, propfans, which are being developed for commercial airplanes. It discusses the three types of propfan engines and the advantages and disadvantages of each. It also tells about the propfan airplanes several companies are developing.

  14. Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan A.

    1997-01-01

    Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

  15. The measurement of electron density in a rocket motor plume

    NASA Astrophysics Data System (ADS)

    Cooper, David A.; Frederick, Robert A.

    1993-06-01

    This paper discusses the development of a diagnostic technique to measure the electron density in a rocket motor plume in order to characterize and rank solid rocket propellants based on their propensity to attenuate the communication signal to a missile. Three techniques were originally investigated as possible low-cost approaches that could be used for plume comparisons as a function of propellant. These approaches consisted of Langmuir probes, electromagnetic coils, and focused microwave probes. The focused microwave probe concept was considered the most appropriate technique to implement for the research to be conducted. The complete design and analysis of a focused microwave probe system operating at 17 GHz was conducted and the selection to determine this operating frequency discussed. Initial estimates of general uncertainty analysis suggest very good results are obtainable using a F-4 lens system and horn diameter of 17 in. for the 17 GHz frequency.

  16. Standard Molded Composite Rocket Pyrogen Igniter - A progress report

    NASA Technical Reports Server (NTRS)

    Lucy, M. H.

    1978-01-01

    The pyrogen igniter has the function to furnish a controlled, high temperature, high pressure gas to ignite solid propellant surfaces in a rocket motor. Present pyrogens consist of numerous inert components. The Standard Molded Pyrogen Igniter (SMPI) consists of three basic parts, a cap with several integrally molded features, an ignition pellet retainer plate, and a tube with additional integrally molded features. A description is presented of an investigation which indicates that the SMPI concept is a viable approach to the design and manufacture of pyrogen igniters for solid propellant rocket motors. For some applications, combining the structural and thermal properties of molded composites can result in the manufacture of lighter assemblies at considerable cost reduction. It is demonstrated that high strength, thin walled tubes with high length to diameter ratios can be fabricated from reinforced plastic molding compound using the displacement compression process.

  17. Direct electrical arc ignition of hybrid rocket motors

    NASA Astrophysics Data System (ADS)

    Judson, Michael I., Jr.

    Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

  18. 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.

  19. Propellant-remaining modeling

    NASA Technical Reports Server (NTRS)

    Torgovitsky, S.

    1991-01-01

    A successful satellite mission is predicted upon the proper maintenance of the spacecraft's orbit and attitude. One requirement for planning and predicting the orbit and attitude is the accurate estimation of the propellant remaining onboard the spacecraft. Focuss is on the three methods that were developed for calculating the propellant budget: the errors associated with each method and the uncertainties in the variables required to determine the propellant remaining that contribute to these errors. Based on these findings, a strategy is developed for improved propellant-remaining estimation. The first method is based on Boyle's law, which related the values of pressure, volume, and temperature (PVT) of an ideal gas. The PVT method is used for the monopropellant and the bipropellant engines. The second method is based on the engine performance tests, which provide data that relate thrust and specific impulse associated with a propellant tank to that tank's pressure. Two curves representing thrust and specific impulse as functions of pressure are then generated using a polynomial fit on the engine performance data. The third method involves a computer simulation of the propellant system. The propellant flow is modeled by creating a conceptual model of the propulsion system configuration, taking into account such factors as the propellant and pressurant tank characteristics, thruster functionality, and piping layout. Finally, a thrust calibration technique is presented that uses differential correction with the computer simulation method of propellant-remaining modeling. Thrust calibration provides a better assessment of thruster performance and therefore enables a more accurate estimation of propellant consumed during a given maneuver.

  20. Launch Vehicle Performance for Bipropellant Propulsion Using Atomic Propellants With Oxygen

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2000-01-01

    Atomic propellants for bipropellant launch vehicles using atomic boron, carbon, and hydrogen were analyzed. The gross liftoff weights (GLOW) and dry masses of the vehicles were estimated, and the 'best' design points for atomic propellants were identified. Engine performance was estimated for a wide range of oxidizer to fuel (O/F) ratios, atom loadings in the solid hydrogen particles, and amounts of helium carrier fluid. Rocket vehicle GLOW was minimized by operating at an O/F ratio of 1.0 to 3.0 for the atomic boron and carbon cases. For the atomic hydrogen cases, a minimum GLOW occurred when using the fuel as a monopropellant (O/F = 0.0). The atomic vehicle dry masses are also presented, and these data exhibit minimum values at the same or similar O/F ratios as those for the vehicle GLOW. A technology assessment of atomic propellants has shown that atomic boron and carbon rocket analyses are considered to be much more near term options than the atomic hydrogen rockets. The technology for storing atomic boron and carbon has shown significant progress, while atomic hydrogen is not able to be stored at the high densities needed for effective propulsion. The GLOW and dry mass data can be used to estimate the cost of future vehicles and their atomic propellant production facilities. The lower the propellant's mass, the lower the overall investment for the specially manufactured atomic propellants.

  1. Solid Rocket Launch Vehicle Explosion Environments

    NASA Technical Reports Server (NTRS)

    Richardson, E. H.; Blackwood, J. M.; Hays, M. J.; Skinner, T.

    2014-01-01

    Empirical explosion data from full scale solid rocket launch vehicle accidents and tests were collected from all available literature from the 1950s to the present. In general data included peak blast overpressure, blast impulse, fragment size, fragment speed, and fragment dispersion. Most propellants were 1.1 explosives but a few were 1.3. Oftentimes the data from a single accident was disjointed and/or missing key aspects. Despite this fact, once the data as a whole was digitized, categorized, and plotted clear trends appeared. Particular emphasis was placed on tests or accidents that would be applicable to scenarios from which a crew might need to escape. Therefore, such tests where a large quantity of high explosive was used to initiate the solid rocket explosion were differentiated. Also, high speed ground impacts or tests used to simulate such were also culled. It was found that the explosions from all accidents and applicable tests could be described using only the pressurized gas energy stored in the chamber at the time of failure. Additionally, fragmentation trends were produced. Only one accident mentioned the elusive "small" propellant fragments, but upon further analysis it was found that these were most likely produced as secondary fragments when larger primary fragments impacted the ground. Finally, a brief discussion of how this data is used in a new launch vehicle explosion model for improving crew/payload survival is presented.

  2. Radiation effect on rocket engine performance

    NASA Technical Reports Server (NTRS)

    Chiu, Huei-Huang

    1988-01-01

    The effects of radiation on the performance of modern rocket propulsion systems operating at high pressure and temperature were recognized as a key issue in the design and operation of various liquid rocket engines of the current and future generations. Critical problem areas of radiation coupled with combustion of bipropellants are assessed and accounted for in the formulation of a universal scaling law incorporated with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and the pertaining data of the Variable Thrust Engine (VTE) and Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low enthalpy engines, such as the VTE, are vulnerable to a substantial performance set back by the radiative loss, whereas the performance of high enthalpy engines such as the SSME, are hardly affected over a broad range of engine operation. Additionally, combustion enhancement by the radiative heating of the propellant has a significant impact in those propellants with high absorptivity. Finally, the areas of research related with radiation phenomena in bipropellant engines are identified.

  3. Capillary device refilling. [liquid rocket propellant tank tests

    NASA Technical Reports Server (NTRS)

    Blatt, M. H.; Merino, F.; Symons, E. P.

    1980-01-01

    An analytical and experimental study was conducted dealing with refilling start baskets (capillary devices) with settled fluid. A computer program was written to include dynamic pressure, screen wicking, multiple-screen barriers, standpipe screens, variable vehicle mass for computing vehicle acceleration, and calculation of tank outflow rate and vapor pullthrough height. An experimental apparatus was fabricated and tested to provide data for correlation with the analytical model; the test program was conducted in normal gravity using a scale-model capillary device and ethanol as the test fluid. The test data correlated with the analytical model; the model is a versatile and apparently accurate tool for predicting start basket refilling under actual mission conditions.

  4. Metallic Hydrogen - Potentially a High Energy Rocket Propellant

    NASA Technical Reports Server (NTRS)

    Cole, John; Silvera, Ike

    2007-01-01

    Pure metallic hydrogen is predicted to have a specific impulse (Isp) of 1700 seconds, but the reaction temperature is too high for current engine materials. Diluting metallic hydrogen with liquid hydrogen can reduce the reaction temperature to levels compatible with current material limits and still provide an Isp greater than 900 s. Metallic hydrogen has not yet been produced on earth, but experimental techniques exist that may change this situation. This paper will provide a brief description of metallic hydrogen and the status of experiments that may soon produce detectable quantities of this material in the lab. Also provided are some characteristics for diluted metallic hydrogen engines and launch vehicles.

  5. Mechanical Slosh Models for Rocket-Propelled Spacecraft

    NASA Technical Reports Server (NTRS)

    Jang, Jiann-Woei; Alaniz, Abram; Yang, Lee; Powers. Joseph; Hall, Charles

    2013-01-01

    Several analytical mechanical slosh models for a cylindrical tank with flat bottom are reviewed. Even though spacecrafts use cylinder shaped tanks, most of those tanks usually have elliptical domes. To extend the application of the analytical models for a cylindrical tank with elliptical domes, the modified slosh parameter models are proposed in this report by mapping an elliptical dome cylindrical tank to a flat top/bottom cylindrical tank while maintaining the equivalent liquid volume. For the low Bond number case, the low-g slosh models were also studied. Those low-g models can be used for Bond number > 10. The current low-g slosh models were also modified to extend their applications for the case that liquid height is smaller than the tank radius. All modified slosh models are implemented in MATLAB m-functions and are collected in the developed MST (Mechanical Slosh Toolbox).

  6. Low-Cost Propellant Launch to LEO from a Tethered Balloon - 'Propulsion Depots' Not 'Propellant Depots'

    NASA Technical Reports Server (NTRS)

    Wilcox, Brian H.; Schneider, Evan G.; Vaughan, David A.; Hall, Jeffrey L.; Yu, Chi Yau

    2011-01-01

    As we have previously reported, it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. Liquid Oxygen/Liquid Hydrogen (LO2/LH2)) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond.

  7. 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.

  8. 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)

  9. 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.

  10. 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)

  11. Nitramine propellants. [gun propellant burning rate

    NASA Technical Reports Server (NTRS)

    Cohen, N. S.; Strand, L. D. (Inventor)

    1978-01-01

    Nitramine propellants without a pressure exponent shift in the burning rate curves are prepared by matching the burning rate of a selected nitramine or combination of nitramines within 10% of burning rate of a plasticized active binder so as to smooth out the break point appearance in the burning rate curve.

  12. Development of a solid propellant viscoelastic dynamic model

    NASA Technical Reports Server (NTRS)

    Hufferd, W. L.; Fitzgerald, J. E.

    1976-01-01

    The results of a one year study to develop a dynamic response model for the Space Shuttle Solid Rocket Motor (SRM) propellant are presented. An extensive literature survey was conducted, from which it was concluded that the only significant variables affecting the dynamic response of the SRM propellant are temperature and frequency. Based on this study, and experimental data on propellants related to the SRM propellant, a dynamic constitutive model was developed in the form of a simple power law with temperature incorporated in the form of a modified power law. A computer program was generated which performs a least-squares curve-fit of laboratory data to determine the model parameters and it calculates dynamic moduli at any desired temperature and frequency. Additional studies investigated dynamic scaling laws and the extent of coupling between the SRM propellant and motor cases. It was found, in agreement with other investigations, that the propellant provides all of the mass and damping characteristics whereas the case provides all of the stiffness.

  13. Navy propeller section characteristics as used in propeller design

    NASA Technical Reports Server (NTRS)

    Weick, Fred E

    1926-01-01

    This report contains artificial aerodynamic characteristics of a set of propeller sections to be used in designing propellers by means of the blade element theory. Characteristics computed from model propeller tests for a single section are extended to cover sections of Navy propeller sections at high Reynolds Number in the variable density tunnel of the NACA.

  14. Microgravity liquid propellant management

    NASA Technical Reports Server (NTRS)

    Hung, R. J.

    1990-01-01

    The requirement to settle or to position liquid fluid over the outlet end of a spacecraft propellant tank prior to main engine restart, poses a microgravity fluid behavior problem. Resettlement or reorientation of liquid propellant can be accomplished by providing optimal acceleration to the spacecraft such that the propellant is reoriented over the tank outlet without any vapor entrainment, any excessive geysering, or any other undersirable fluid motion for the space fluid management under microgravity environment. The most efficient technique is studied for propellant resettling through the minimization of propellant usage and weight penalties. Both full scale and subscale liquid propellant tank of Space Transfer Vehicle were used to simulate flow profiles for liquid hydrogen reorientation over the tank outlet. In subscale simulation, both constant and impulsive resettling acceleration were used to simulate the liquid flow reorientation. Comparisons between the constant reverse gravity acceleration and impulsive reverse gravity acceleration to be used for activation of propellant resettlement shows that impulsive reverse gravity thrust is superior to constant reverse gravity thrust.

  15. Autonomous production of propellants

    NASA Technical Reports Server (NTRS)

    Ramohalli, Kumar; Schallhorn, P. A.

    1990-01-01

    The autonomous production of propellants is addressed. Since 80 to 90 percent of a spacecraft's mass is typically propellants, it is advantageous to produce propellants in strategic locations en route to, and at, the desired mission destination. This reduces the weight of the spacecraft and the cost of each mission. Since one of the primary goals of the space program is safety, a totally automated propellant production system is desirable. This system would remove, from hostile, high-risk extraterrestrial environments, the constant human intervention currently required in the production of many propellants. This enables the exploration of space to be more than the search for and production of fuel. As a proof-of-concept demonstration, one specific case was chosen for this study. That case was a composite propellant processor (the principle is more important than the application), and the specific processor used saved SERC the considerable cost of acquiring a new liquid propellant processor that would also have required similar automation.

  16. Nitramine smokeless propellant research

    NASA Technical Reports Server (NTRS)

    Cohen, N. S.; Strand, L. P.

    1977-01-01

    A transient ballistics and combustion model is derived to represent the closed vessel experiment that is widely used to characterize propellants. A computer program is developed to solve the time-dependent equations, and is applied to explain aspects of closed vessel behavior. In the case of nitramine propellants the cratering of the burning surface associated with combustion above break-point pressures augments the effective burning rate as deduced from the closed vessel experiment. Low pressure combustion is significantly affected by the ignition process and, in the case of nitramine propellants, by the developing and changing surface structure. Thus, burning rates deduced from the closed vessel experiment may or may not agree with those measured in the equilibrium strand burner. Series of T burner experiments are performed to compare the combustion instability characteristics of nitramine (HMX) containing propellants and ammonium perchlorate (AP)propellants. Although ash produced by more fuel rich propellants could have provided mechanical suppression, results from clean-burning propellants permit the conclusion that HMX reduces the acoustic driving.

  17. Ricardo Dyrgalla (1910-1970), pioneer of rocket development in Argentina

    NASA Astrophysics Data System (ADS)

    de León, Pablo

    2009-12-01

    One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

  18. XDT in HTPB propellant from steel flyer plate impact tests

    NASA Astrophysics Data System (ADS)

    Tanaka, Katsumi; Noda, Keiichiro; Hyodo, Yukio; Nakamura, Hiroyuki; Kosaka, Katsuaki; Nakayama, Takashi; Katayama, Masahide; Takeba, Atsushi

    2000-04-01

    Several experiments simulating ground impact explosion following the command destruction of a launch vehicle have been performed using HTPB propellant samples of mass 460 to 940 kg impacted by a steel flyer plate. Impact velocities were varied from 135 m/s to 170 m/s. Strong explosions were observed at impact velocities higher than 150 m/s for tests of solid rocket propellant weighting 460 kg. The XDT (Unknown to Detonation Transition) is studied using a bulk failure reaction model including strain rate effect. Computational results are compared with observed blast waves for various impact velocities. The present model has been successfully applied to 22 inch Critical Diameter tests for SRMU HTPB propellant.

  19. Effect of inert propellant injection on Mars ascent vehicle performance

    NASA Technical Reports Server (NTRS)

    Colvin, James E.; Landis, Geoffrey A.

    1992-01-01

    A Mars ascent vehicle is limited in performance by the propellant which can be brought from Earth. In some cases the vehicle performance can be improved by injecting inert gas into the engine, if the inert gas is available as an in-situ resource and does not have to be brought from Earth. Carbon dioxide, nitrogen, and argon are constituents of the Martian atmosphere which could be separated by compressing the atmosphere, without any chemical processing step. The effect of inert gas injection on rocket engine performance was analyzed with a numerical combustion code that calculated chemical equilibrium for engines of varying combustion chamber pressure, expansion ratio, oxidizer/fuel ratio, and inert injection fraction. Results of this analysis were applied to several candidate missions to determine how the required mass of return propellant needed in low Earth orbit could be decreased using inert propellant injection.

  20. Characterization of aluminum/RP-1 gel propellant properties

    NASA Technical Reports Server (NTRS)

    Rapp, Douglas C.; Zurawski, Robert L.

    1988-01-01

    Research efforts are being conducted by the NASA Lewis Research Center to formulate and characterize the properties of Al/RP-1 and RP-1 gelled propellants for rocket propulsion systems. Twenty four different compositions of gelled fuels were formulated with 5 and 16 micron, atomized aluminum powder in RP-1. The total solids concentration in the propellant varied from 5 to 60 wt percent. Tests were conducted to evaluate the stability and rheological characteristics of the fuels. Physical separation of the solids occurred in fuels with less than 50 wt percent solids concentration. The rheological characteristics of the Al/RP-1 fuels varied with solids concentration. Both thixotropic and rheopectic gel behavior were observed. The unmetallized RP-1 gels, which were formulated by a different technique than the Al/RP-1 gels, were highly viscoelastic. A history of research efforts which were conducted to formulate and characterize the properties of metallized propellants for various applications is also given.

  1. Composite propellant tank study for very low cost space transportation

    NASA Technical Reports Server (NTRS)

    Moser, D. J.; Keith, E. L.

    1992-01-01

    A study of life-cycle cost is conducted to determine acceptable options for composite propellant tanks at low cost and weight and for use at moderate pressures. The review examines all cost issues relevant to the production, mass, applications, and reliability of the tanks for pressure-fed rockets. Specific attention is given to the manufacturing and life-cycle issues relevant to the use of composite materials in this application since composites are effective materials for liquid propellant tanks. Specific costs and parametric considerations are given for several tank candidates with 62,303-lb capacities. The mass sensitivity of the fourth stage for the concept vehicle is shown to be high, and the use of a 325-psi fourth-stage tank is shown to yield the minimum cost/lb for the stage. Wound S-glass/epoxy composites can be employed as cost-effective replacements for steel in the design of liquid-propellant tanks.

  2. Laboratory test methods for combustion stability properties of solid propellants

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Brown, R. S.

    1992-01-01

    An overview is presented of experimental methods for determining the combustion-stability properties of solid propellants. The methods are generally based on either the temporal response to an initial disturbance or on external methods for generating the required oscillations. The size distribution of condensed-phase combustion products are characterized by means of the experimental approaches. The 'T-burner' approach is shown to assist in the derivation of pressure-coupled driving contributions and particle damping in solid-propellant rocket motors. Other techniques examined include the rotating-valve apparatus, the impedance tube, the modulated throat-acoustic damping burner, and the magnetic flowmeter. The paper shows that experimental methods do not exist for measuring the interactions between acoustic velocity oscillations and burning propellant.

  3. 'RCHX-1-STORM' first Slovenian meteorological rocket program

    NASA Astrophysics Data System (ADS)

    Kerstein, Aleksander; Matko, Drago; Trauner, Amalija; Britovšek, Zvone

    2004-08-01

    Astronautic and Rocket Society Celje (ARSC) formed a special working team for research and development of a small meteorological hail suppression rocket in the 70th. The hail suppression system was established in former Yugoslavia in the late 60th as an attempt to protect important agricultural regions from one of the summer's most vicious storm. In this time Slovenia was a part of Yugoslavia as one of the federal republic with relative high developed agricultural region production. The Rocket program 'RCHX-STORM' was a second attempt, for Slovenia indigenously developed in the production of meteorological hail suppression rocket. ARSC has designed a family of small sounding rocket that were based on highly promising hybrid propellant propulsion. Hybrid propulsion was selected for this family because it was offering low cost, save production and operation and simple logistics. Conventional sounding rockets use solid propellant motor for their propulsion. The introduction of hybrid motors has enabled a considerable decrease in overall cost. The transportation handling and storage procedures were greatly simplified due to the fact that a hybrid motor was not considered as explosive matter. A hybrid motor may also be designed to stand a severe environment without resorting to conditioning arrangements. The program started in the late 70th when the team ARSC was integrated in the Research and Development Institute in Celje (RDIC). The development program aimed to produce three types of meteorological rockets with diameters 76, 120 and 160 mm. Development of the RCHX-76 engine and rocket vehicle including flight certification has been undertaken by a joint team comprising of the ARCS, RDIC and the company Cestno podjetje Celje (CPC), Road building company Celje. Many new techniques and methods were used in this program such as computer simulation of external and internal ballistics, composite materials for rocket construction, intensive static testing of models and

  4. Russian Meteorological and Geophysical Rockets of New Generation

    NASA Astrophysics Data System (ADS)

    Yushkov, V.; Gvozdev, Yu.; Lykov, A.; Shershakov, V.; Ivanov, V.; Pozin, A.; Afanasenkov, A.; Savenkov, Yu.; Kuznetsov, V.

    2015-09-01

    To study the process in the middle and upper atmosphere, ionosphere and near-Earth space, as well as to monitor the geophysical environment in Russian Federal Service for Hydrology and Environmental Monitoring (ROSHYDROMET) the development of new generation of meteorological and geophysical rockets has been completed. The modern geophysical research rocket system MR-30 was created in Research and Production Association RPA "Typhoon". The basis of the complex MR-30 is a new geophysical sounding rocket MN-300 with solid propellant, Rocket launch takes place at an angle of 70º to 90º from the launcher, which is a farm with a guide rail type required for imparting initial rotation rocket. The Rocket is spin stabilized with a spin rate between 5 and 7 Hz. Launch weight is 1564 kg, and the mass of the payload of 50 to 150 kg. MR-300 is capable of lifting up to 300 km, while the area of dispersion points for booster falling is an ellipse with parameters 37x 60 km. The payload of the rocket MN-300 consists of two sections: a sealed, located below the instrument compartment, and not sealed, under the fairing. Block of scientific equipment is formed on the platform in a modular layout. This makes it possible to solve a wide range of tasks and conduct research and testing technologies using a unique environment of space, as well as to conduct technological experiments testing and research systems and spacecraft equipment. New Russian rocket system MERA (MEteorological Rocket for Atmospheric Research) belongs to so called "dart" technique that provide lifting of small scientific payload up to altitude 100 km and descending with parachute. It was developed at Central Aerological Observatory jointly with State Unitary Enterprise Instrument Design Bureau. The booster provides a very rapid acceleration to about Mach 5. After the burning phase of the buster the dart is separated and continues ballistic flight for about 2 minutes. The dart carries the instrument payload+ parachute

  5. Parallelization of Rocket Engine System Software (Press)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1996-01-01

    The main goal is to assess parallelization requirements for the Rocket Engine Numeric Simulator (RENS) project which, aside from gathering information on liquid-propelled rocket engines and setting forth requirements, involve a large FORTRAN based package at NASA Lewis Research Center and TDK software developed by SUBR/UWF. The ultimate aim is to develop, test, integrate, and suitably deploy a family of software packages on various aspects and facets of rocket engines using liquid-propellants. At present, all project efforts by the funding agency, NASA Lewis Research Center, and the HBCU participants are disseminated over the internet using world wide web home pages. Considering obviously expensive methods of actual field trails, the benefits of software simulators are potentially enormous. When realized, these benefits will be analogous to those provided by numerous CAD/CAM packages and flight-training simulators. According to the overall task assignments, Hampton University's role is to collect all available software, place them in a common format, assess and evaluate, define interfaces, and provide integration. Most importantly, the HU's mission is to see to it that the real-time performance is assured. This involves source code translations, porting, and distribution. The porting will be done in two phases: First, place all software on Cray XMP platform using FORTRAN. After testing and evaluation on the Cray X-MP, the code will be translated to C + + and ported to the parallel nCUBE platform. At present, we are evaluating another option of distributed processing over local area networks using Sun NFS, Ethernet, TCP/IP. Considering the heterogeneous nature of the present software (e.g., first started as an expert system using LISP machines) which now involve FORTRAN code, the effort is expected to be quite challenging.

  6. Automated Propellant Blending

    NASA Technical Reports Server (NTRS)

    Hohmann, Carl W. (Inventor); Harrington, Douglas W. (Inventor); Dutton, Maureen L. (Inventor); Tipton, Billy Charles, Jr. (Inventor); Bacak, James W. (Inventor); Salazar, Frank (Inventor)

    2000-01-01

    An automated propellant blending apparatus and method that uses closely metered addition of countersolvent to a binder solution with propellant particles dispersed therein to precisely control binder precipitation and particle aggregation is discussed. A profile of binder precipitation versus countersolvent-solvent ratio is established empirically and used in a computer algorithm to establish countersolvent addition parameters near the cloud point for controlling the transition of properties of the binder during agglomeration and finishing of the propellant composition particles. The system is remotely operated by computer for safety, reliability and improved product properties, and also increases product output.

  7. Automated Propellant Blending

    NASA Technical Reports Server (NTRS)

    Hohmann, Carl W. (Inventor); Harrington, Douglas W. (Inventor); Dutton, Maureen L. (Inventor); Tipton, Billy Charles, Jr. (Inventor); Bacak, James W. (Inventor); Salazar, Frank (Inventor)

    1999-01-01

    An automated propellant blending apparatus and method uses closely metered addition of countersolvent to a binder solution with propellant particles dispersed therein to precisely control binder precipitation and particle aggregation. A profile of binder precipitation versus countersolvent-solvent ratio is established empirically and used in a computer algorithm to establish countersolvent addition parameters near the cloud point for controlling the transition of properties of the binder during agglomeration and finishing of the propellant composition particles. The system is remotely operated by computer for safety, reliability and improved product properties, and also increases product output.

  8. High temperature propellant development

    NASA Technical Reports Server (NTRS)

    Anderson, F. A.

    1981-01-01

    It is reported that the neccessary technology has been developed and demonstrated for the manufacture of heat-sterilizable solid propellants which meet specific ballistic goals. It is shown that: (1) phosphate doping of ammonium perchlorate significantly enhances the thermal stability of the substance; (2) grinding the ammonium perchlorate to reduce particle size further increases thermal stability; and (3) unsaturated polymers such as the polybutadienes can be successfully used in a heat-sterilizable propellant system. Among the topics considered by the study are oxidizers, dopants, binders, and the thermal cycling of 70 lb and 600 lb propellant grains.

  9. High Energy Density Matter for Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Carrick, Patrick G.

    1996-01-01

    The objective of the High Energy Density Matter (HEDM) program is to identify, develop, and exploit high energy atomic and molecular systems as energetic sources for rocket propulsion applications. It is a high risk, high payoff program that incorporates both basic and applied research, experimental and theoretical efforts, and science and engineering efforts. The HEDM program is co-sponsored by the Air Force Office of Scientific Research (AFOSR) and the Phillips Laboratory (PURKS). It includes both in-house and contracted University/Industry efforts. Technology developed by the HEDM program offers the opportunity for significant breakthroughs in propulsion system capabilities over the current state-of-the-art. One area of great interest is the use of cryogenic solids to increase the density of the propellant and to act as a stable matrix for storage of energetic materials. No cryogenic solid propellant has ever been used in a rocket, and there remain engineering challenges to such a propellant. However, these solids would enable a wide class of highly energetic materials by providing an environment that is at very low temperatures and is a physical barrier to recombination or energy loss reactions. Previous to our experiments only hydrogen atoms had been isolated in solid hydrogen. To date we have succeeded in trapping B, Al, Li, N, and Mg atoms in solid H2. Small molecules, such as B2 and LiB, are also of interest. Current efforts involve the search for new energetic small molecules, increasing free radical concentrations up to 5 mole percent, and scale-up for propulsion testing.

  10. Turbo Pump Fed Micro-Rocket Engine

    NASA Astrophysics Data System (ADS)

    Miotti, P.; Tajmar, M.; Seco, F.; Guraya, C.; Perennes, F.; Soldati, A.; Lang, M.

    2004-10-01

    Micro-satellites (from 10kg up to 100kg) have mass, volume, and electrical power constraints due to their low dimensions. These limitations lead to the lack in currently available active orbit control systems in micro-satellites. Therefore, a micro-propulsion system with a high thrust to mass ratio is required to increase the potential functionality of small satellites. Mechatronic is presently working on a liquid bipropellant micro-rocket engine under contract with ESA (Contract No.16914/NL/Sfe - Micro-turbo-machinery Based Bipropellant System Using MNT). The advances in Mechatronic's project are to realise a micro-rocket engine with propellants pressurised by micro-pumps. The energy for driving the pumps would be extracted from a micro-turbine. Cooling channels around the nozzle would be also used in order to maintain the wall material below its maximum operating temperature. A mass budget comparison with more traditional pressure-fed micro-rockets shows a real benefit from this system in terms of mass reduction. In the paper, an overview of the project status in Mechatronic is presented.

  11. Key technology for reusable rocket engine turbopump

    NASA Astrophysics Data System (ADS)

    Okayasu, A.; Ohta, T.; Kamijyo, A.; Yamada, H.

    2002-03-01

    Recently, there has been an increased need for evolved space transportation and the research of reusable rocket which enable low cost and high reliability and is generating a lot of interest all over the world. In the USA, the development of reusable launch vehicle "Venture Star" which will be used instead of space shuttle is planned and its half scale model "X-33" was developed for the first flight in 1999. In Japan, there has been agreement on the main points to develop the rocket type RLV based on the technology of H-IIA, HOPE-X before developing space plane type RLV. The planned reusable rocket engine was LOX/LH2 as propellant, has 100- 200 ton thrust and has a throttling capability. In addition, long life and high reliability are required for the engine system including LOX/LH2 turbopump. The paper introduces some key technologies for the reusable turbopump which IHI is promoting for research and development with NAL and Tohoku University.

  12. Welded Titanium Case for Space-Probe Rocket Motor

    NASA Technical Reports Server (NTRS)

    Brothers, A. J.; Boundy, R. A.; Martens, H. E.; Jaffe, L. D.

    1959-01-01

    The high strength-to-weight ratio of titanium alloys suggests their use for solid-propellant rocket-motor cases for high-performance orbiting or space-probe vehicles. The paper describes the fabrication of a 6-in.-diam., 0.025-in.-wall rocket-motor from the 6A1-4V titanium alloy. The rocket-motor case, used in the fourth stage of a successful JPL-NASA lunar-probe flight, was constructed using a design previously proven satisfactory for Type 410 stainless steel. The nature and scope of the problems peculiar to the use of the titanium alloy, which effected an average weight saving of 34%, are described.

  13. Radiation/convection coupling in rocket motors and plumes

    NASA Technical Reports Server (NTRS)

    Farmer, R. C.; Saladino, A. J.

    1993-01-01

    The three commonly used propellant systems - H2/O2, RP-1/O2, and solid propellants - primarily radiate as molecular emitters, non-scattering small particles, and scattering larger particles, respectively. Present technology has accepted the uncoupling of the radiation analysis from that of the flowfield. This approximation becomes increasingly inaccurate as one considers plumes, interior rocket chambers, and nuclear rocket propulsion devices. This study will develop a hierarchy of methods which will address radiation/convection coupling in all of the aforementioned propulsion systems. The nature of the radiation/convection coupled problem is that the divergence of the radiative heat flux must be included in the energy equation and that the local, volume-averaged intensity of the radiation must be determined by a solution of the radiative transfer equation (RTE). The intensity is approximated by solving the RTE along several lines of sight (LOS) for each point in the flowfield. Such a procedure is extremely costly; therefore, further approximations are needed. Modified differential approximations are being developed for this purpose. It is not obvious which order of approximations are required for a given rocket motor analysis. Therefore, LOS calculations have been made for typical rocket motor operating conditions in order to select the type approximations required. The results of these radiation calculations, and the interpretation of these intensity predictions are presented herein.

  14. 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.

  15. 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…

  16. Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

    NASA Astrophysics Data System (ADS)

    Forde, Scott; Bulman, Mel; Neill, Todd

    2006-07-01

    Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

  17. 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.

  18. Effect of Cumulative Damage on Rocket Motor Service Life

    NASA Astrophysics Data System (ADS)

    Gligorijević, Nikola; Živković, Saša; Subotić, Sredoje; Rodić, Vesna; Gligorijević, Ivan

    2015-10-01

    Two series of antihail rocket propellant grains failed only 3 months after production, due to the appearance of cracks in the grain channel. Structural integrity analysis demonstrated sufficient reliability at the beginning of service life. Further analysis showed that under temperature loads, cumulative damage during the short period in field stocks caused the grain failure, despite the established opinion that such failure can become significant only after lengthy storage. A linear cumulative damage law is evaluated by exposing a number of hydroxyl-terminated polybutadiene (HTPB) composite propellant specimens to different but constant stress levels. The analysis showed that cumulative damage must not be overlooked at the design stage. Further, a positive correlation between the propellant cumulative damage law and tensile strength is strongly indicated.

  19. 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.

  20. Nitramine smokeless propellant research

    NASA Technical Reports Server (NTRS)

    1977-01-01

    A transient ballistics and combustion model was derived to represent the closed vessel experiment that is widely used to characterize propellants. The model incorporates the nitramine combustion mechanisms. A computer program was developed to solve the time dependent equations, and was applied to explain aspects of closed vessel behavior. It is found that the rate of pressurization in the closed vessel is insufficient at pressures of interest to augment the burning rate by time dependent processes. Series of T-burner experiments were performed to compare the combustion instability characteristics of nitramine (HMX) containing propellants and ammonium perchlorate (AP) propellants. It is found that the inclusion of HMX consistently renders the propellant more stable.

  1. NASA propeller noise research

    NASA Technical Reports Server (NTRS)

    Greene, G. C.

    1980-01-01

    The research in propeller noise prediction, noise/performance optimization, and interior reduction is described. Selected results are presented to illustrate the status of the technology and the direction of future research.

  2. Spinning Rocket Simulator Turntable Design

    NASA Technical Reports Server (NTRS)

    Miles, Robert W.

    2001-01-01

    Contained herein is the research and data acquired from the Turntable Design portion of the Spinning Rocket Simulator (SRS) project. The SRS Project studies and eliminates the effect of coning on thrust-propelled spacecraft. This design and construction of the turntable adds a structural support for the SRS model and two degrees of freedom. The two degrees of freedom, radial and circumferential, will help develop a simulated thrust force perpendicular to the plane of the spacecraft model while undergoing an unstable coning motion. The Turntable consists of a ten-foot linear track mounted to a sprocket and press-fit to a thrust bearing. A two-inch high column grounded by a Triangular Baseplate supports this bearing and houses the slip rings and pressurized, air-line swivel. The thrust bearing allows the entire system to rotate under the moment applied through the chain-driven sprocket producing a circumferential degree of freedom. The radial degree of freedom is given to the model through the helically threaded linear track. This track allows the Model Support and Counter Balance to simultaneously reposition according to the coning motion of the Model. Two design factors that hinder the linear track are bending and twist due to torsion. A Standard Aluminum "C" channel significantly reduces these two deflections. Safety considerations dictate the design of all the components involved in this project.

  3. NASA's Advanced solid rocket motor

    NASA Astrophysics Data System (ADS)

    Mitchell, Royce E.

    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  4. NASA's Advanced solid rocket motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1993-01-01

    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  5. Two stage turbine for rockets

    NASA Astrophysics Data System (ADS)

    Veres, Joseph P.

    1993-11-01

    The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

  6. Two stage turbine for rockets

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.

    1993-01-01

    The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

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

    NASA Technical Reports Server (NTRS)

    Thorpe, Douglas G.

    1991-01-01

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

  8. Liquid Propulsion: Propellant Feed System Design. Chapter 2.3.11

    NASA Technical Reports Server (NTRS)

    Cannon, James L.

    2010-01-01

    The propellant feed system of a liquid rocket engine determines how the propellants are delivered from the tanks to the thrust chamber. They are generally classified as either pressure fed or pump fed. The pressure-fed system is simple and relies on tank pressures to feed the propellants into the thrust chamber. This type of system is typically used for space propulsion applications and auxiliary propulsion applications requiring low system pressures and small quantities of propellants. In contrast, the pump-fed system is used for high pressure, high performance applications. The selection of one propellant feed system over another is determined based on design trade studies at both the engine and vehicle levels. This chapter first provides a brief overview of the basic configurations of pressure-fed systems. Pump-fed systems are then discussed with greater detail given to the turbomachinery design. Selected design requirements and configurations are provided.

  9. Thermodynamic Vent System Applied as Propellant Delivery System for Air Force

    NASA Technical Reports Server (NTRS)

    1996-01-01

    Responding to a request from the Air Force, NASA Lewis Research Center engineers designed a combination pressure control and propellant delivery system based on thermodynamic vent system (TVS) technology. The Air Force is designing a new type of orbit transfer vehicle that uses energy from sunlight to both propel and power the vehicle. Because this vehicle uses propellant at a substantially slower rate than higher-energy rockets, it needed the Lewis-developed TVS technology for long-duration storage of cryogen propellants. Lewis engineers, in conjunction with industry partners, showed how this TVS technology could also be used to deliver propellant to the thruster. The Air Force has now begun the ground test demonstration phase. After successful completion of ground testing, the Air Force plans to use this technology in a space flight as early as 1999.

  10. The aerodynamics of propellers

    NASA Astrophysics Data System (ADS)

    Wald, Quentin R.

    2006-02-01

    The theory and the design of propellers of minimum induced loss is treated. The pioneer analysis of this problem was presented more than half a century ago by Theodorsen, but obscurities in his treatment and inaccuracies and limited coverage in his tables of the Goldstein circulation function for helicoidal vortex sheets have not been remedied until the present work which clarifies and extends his work. The inverse problem, the prediction of the performance of a given propeller of arbitrary form, is also treated. The theory of propellers of minimum energy loss is dependent on considerations of a regular helicoidal trailing vortex sheet; consequently, a more detailed discussion of the dynamics of vortex sheets and the consequences of their instability and roll up is presented than is usually found in treatments of propeller aerodynamics. Complete and accurate tables of the circulation function are presented. Interference effects between a fuselage or a nacelle and the propeller are considered. The regimes of propeller, vortex ring, and windmill operation are characterized.

  11. Advanced propeller research

    NASA Technical Reports Server (NTRS)

    Groeneweg, John F.; Bober, Lawrence J.

    1987-01-01

    Resent results of aerodynamic and acoustic research on both single and counter-rotation propellers are reviewed. Data and analytical results are presented for three propellers: SR-7A, the single rotation design used in the NASA Propfan Test Assessment (PTA); and F7-A7, the 8+8 counterrotating design used in the proof-of-concept Unducted Fan (UDF) engine. In addition to propeller efficiencies, cruise and takeoff noise, and blade pressure data, off-design phenomena involving formation of leading edge vortices are described. Aerodynamic and acoustic computational results derived from three-dimensional Euler and acoustic radiation codes are presented. Research on unsteady flows, which are particularly important for understanding counterrotation interaction noise, unsteady loading effects on acoustics, and flutter or forced response is described. The first results of three-dimensional unsteady Euler solutions are illustrated for a single rotation propeller at an angle of attack and for a counterrotation propeller. Basic experimental and theoretical results from studies of the unsteady aerodynamics of oscillating cascades are outlined. Finally, advanced concepts involving swirl recovery vanes and ultra bypass ducted propellers are discussed.

  12. Propellant Readiness Level: A Methodological Approach to Propellant Characterization

    NASA Technical Reports Server (NTRS)

    Bossard, John A.; Rhys, Noah O.

    2010-01-01

    A methodological approach to defining propellant characterization is presented. The method is based on the well-established Technology Readiness Level nomenclature. This approach establishes the Propellant Readiness Level as a metric for ascertaining the readiness of a propellant or a propellant combination by evaluating the following set of propellant characteristics: thermodynamic data, toxicity, applications, combustion data, heat transfer data, material compatibility, analytical prediction modeling, injector/chamber geometry, pressurization, ignition, combustion stability, system storability, qualification testing, and flight capability. The methodology is meant to be applicable to all propellants or propellant combinations; liquid, solid, and gaseous propellants as well as monopropellants and propellant combinations are equally served. The functionality of the proposed approach is tested through the evaluation and comparison of an example set of hydrocarbon fuels.

  13. The influence of bonding agents in improving interactions in composite propellants determined using image analysis.

    PubMed

    Dostanić, J; Husović, T V; Usćumlić, G; Heinemann, R J; Mijin, D

    2008-12-01

    Binder-oxidizer interactions in rocket composite propellants can be improved using adequate bonding agents. In the present work, the effectiveness of different 1,3,5-trisubstituted isocyanurates was determined by stereo and metallographic microscopy and using the software package Image-Pro Plus. The chemical analysis of samples was performed by a scanning electron microscope equipped for energy dispersive spectrometry. PMID:19094035

  14. 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.

  15. 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.

  16. 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...

  17. 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.

  18. 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)

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. A Practical, Affordable Cryogenic Propellant Depot Based on ULA's Flight Experience

    NASA Technical Reports Server (NTRS)

    Kutter, Bernard F.; Zegler, Frank; O'Neil, Gary; Pitchford, Brian

    2008-01-01

    Mankind is embarking on the next step in the journey of human exploration. We are returning to the moon and eventually moving to Mars and beyond. The current Exploration architecture seeks a balance between the need for a robust infrastructure on the lunar surface, and the performance limitations of Ares I and V. The ability to refuel or top-off propellant tanks from orbital propellant depots offers NASA the opportunity to cost effectively and reliably satisfy these opposing requirements. The ability to cache large orbital quantities of propellant is also an enabling capability for missions to Mars and beyond. This paper describes an option for a propellant depot that enables orbital refueling supporting Exploration, national security, science and other space endeavors. This proposed concept is launched using a single EELV medium class rocket and thus does not require any orbital assembly. The propellant depot provides cryogenic propellant storage that utilizes flight proven technologies augmented with technologies currently under development. The propellant depot system, propellant management, flight experience, and key technologies are also discussed. Options for refueling the propellant depot along with an overview of Exploration architecture impacts are also presented.

  4. Rocket pioneer Robert Goddard: A micro-biography (pt 3/3)

    NASA Technical Reports Server (NTRS)

    1970-01-01

    Archive footage of Robert Goddard, rocket pioneer of the 1920's and '30's. Ahead of his time, and the first to use liquid propellant. From the 'Moonwalk Series: Episode 1 - 'The Day Before''. A four part documentary series made in the 1970's about the Apollo 11 mission.

  5. Rocket pioneer Robert Goddard: A micro-biography (pt 2/3)

    NASA Technical Reports Server (NTRS)

    1970-01-01

    Archive footage of Robert Goddard, rocket pioneer of the 1920's and '30's. Ahead of his time, and the first to use liquid propellant. From the 'Moonwalk Series: Episode 1 - 'The Day Before''. A four part documentary series made in the 1970's about the Apollo 11 mission.

  6. Rocket pioneer Robert Goddard: A micro-biography (pt 1/3)

    NASA Technical Reports Server (NTRS)

    1970-01-01

    Archive footage of Robert Goddard, rocket pioneer of the 1920's and '30's. Ahead of his time, and the first to use liquid propellant. From the 'Moonwalk Series: Episode 1 - 'The Day Before''. A four part documentary series made in the 1970's about the Apollo 11 mission.

  7. Study of solid rocket motor for space shuttle booster, volume 2, book 3, appendix A

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A systems requirements analysis for the solid propellant rocket engine to be used with the space shuttle was conducted. The systems analysis was developed to define the physical and functional requirements for the systems and subsystems. The operations analysis was performed to identify the requirements of the various launch operations, mission operations, ground operations, and logistic and flight support concepts.

  8. Study of solid rocket motor for space shuttle booster, Volume 3: Program acquisition planning

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The program planning acquisition functions for the development of the solid propellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.

  9. Combustion characteristics of SMX and SMX based propellants

    NASA Astrophysics Data System (ADS)

    Reese, David A.

    This work investigates the combustion of the new solid nitrate ester 2,3-hydroxymethyl-2,3-dinitro-1,4-butanediol tetranitrate (SMX, C6H 8N6O16). SMX was synthesized for the first time in 2008. It has a melting point of 85 °C and oxygen balance of 0% to CO 2, allowing it to be used as an energetic additive or oxidizer in solid propellants. In addition to its neat combustion characteristics, this work also explores the use of SMX as a potential replacement for nitroglycerin (NG) in double base gun propellants and as a replacement for ammonium perchlorate in composite rocket propellants. The physical properties, sensitivity characteristics, and combustion behaviors of neat SMX were investigated. Its combustion is stable at pressures of up to at least 27.5 MPa (n = 0.81). The observed flame structure is nearly identical to that of other double base propellant ingredients, with a primary flame attached at the surface, a thick isothermal dark zone, and a luminous secondary flame wherein final recombination reactions occur. As a result, the burning rate and primary flame structure can be modeled using existing one-dimensional steady state techniques. A zero gas-phase activation energy approximation results in a good fit between modeled and observed behavior. Additionally, SMX was considered as a replacement for nitroglycerin in a double base propellant. Thermochemical calculations indicate improved performance when compared with the common double base propellant JA2 at SMX loadings above 40 wt-%. Also, since SMX is a room temperature solid, migration may be avoided. Like other nitrate esters, SMX is susceptible to decomposition over long-term storage due to the presence of excess acid in the crystals; the addition of stabilizers (e.g., derivatives of urea) during synthesis should be sufficient to prevent this. the addition of Both unplasticized and plasticized propellants were formulated. Thermal analysis of unplasticized propellant showed a distinct melt

  10. A Plasma Rocket Demonstration on the International Space Station

    NASA Astrophysics Data System (ADS)

    Petro, A.

    2002-01-01

    in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One feature of this concept is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. For this reason the system is called the Variable Specific Impulse Magneto-plasma Rocket or VASIMR. This ability to vary specific impulse and thrust will allow for optimum low thrust interplanetary trajectories and results in shorter trip times than is possible with fixed specific impulse systems while preserving adequate payload margins. demonstrations are envisioned. A ground-based experiment of a low-power VASIMR prototype rocket is currently underway at the Advanced Space Propulsion Laboratory. The next step is a proposal to build and fly a 25-kilowatt VASIMR rocket as an external payload on the International Space Station. This experiment will provide an opportunity to demonstrate the performance of the rocket in space and measure the induced environment. The experiment will also utilize the space station for its intended purpose as a laboratory with vacuum conditions that cannot be matched by any laboratory on Earth. propulsion on the space station. An electric propulsion system like VASIMR, if provided with sufficient electrical power, could provide continuous drag force compensation for the space station. Drag compensation would eliminate the need for reboosting the station, an operation that will consume about 60 metric tons of propellant in a ten-year period. In contrast, an electric propulsion system would require very little propellant. In fact, a system like VASIMR can use waste hydrogen from the station's life support system as its propellant. This waste hydrogen is otherwise dumped overboard. Continuous drag compensation would also improve the microgravity conditions on the station. So electric propulsion can reduce propellant

  11. Technical prospects for utilizing extraterrestrial propellants for space exploration

    NASA Technical Reports Server (NTRS)

    Linne, Diane L.; Meyer, Michael L.

    1991-01-01

    NASA's LeRC has supported several efforts to understand how lunar and Martian produced propellants can be used to their best advantage for space exploration propulsion. A discussion of these efforts and their results is presented. A Manned Mars Mission Analysis Study identified that a more thorough technology base for propellant production is required before the the net economic benefits of in situ propellants can be determined. Evaluation of the materials available on the moon indicated metal/oxygen combinations are the most promising lunar propellants. A hazard analysis determined that several lunar metal/LOX monopropellants could be safely worked with in small quantities, and a characterization study was initiated to determine the physical and chemical properties of potential lunar monopropellant formulations. A bipropellant metal/oxygen subscale test engine which utilizes pneumatic injection of powdered metal is being pursued as an alternative to the monopropellant systems. The technology for utilizing carbon monoxide/oxygen, a potential Martian propellant, was studied in subscale ignition and rocket performance experiments.

  12. Hydrazine engine plume contamination mapping. [measuring instruments for rocket exhaust from liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Chirivella, J. E.

    1975-01-01

    Instrumentation for the measurement of plume exhaust specie deposition rates were developed and demonstrated. The instruments, two sets of quartz crystal microbalances, were designed for low temperature operation in the back flow and variable temperature operation in the core flow regions of an exhaust plume. These quartz crystal microbalances performed nominally, and measurements of exhaust specie deposition rates for 8400 number of pulses for a 0.1-lb monopropellant thruster are reported.

  13. Advanced propeller research

    NASA Technical Reports Server (NTRS)

    Groeneweg, John F.; Bober, Lawrence J.

    1990-01-01

    Recent results of aerodynamic and acoustic research on both single rotation and counterrotation propellers are reviewed. Data and analytical results are presented for three propellers: SR-7A, the single rotation design used in the NASA Propfan Test Assessment (PTA) flight program; CRP-X1, the initial 5+5 Hamilton Standard counterrotating design; and F7-A7, the 8+8 counterrotating G.E. design used in the proof of concept Unducted Fan (UDF) engine. In addition to propeller efficiencies, cruise and takeoff noise, and blade pressure data, off-design phenomena involving formation of leading edge vortexes are described. Aerodynamic and acoustic computational results derived from 3-D Euler and acoustic radiation codes are presented. Research on unsteady flows which are particularly important for understanding counterrotation interaction noise, unsteady loading effects on acoustics, and flutter or forced response is described. The first results of 3-D unsteady Euler solutions are illustrated for a single rotation propeller at angle of attack and for a counterrotation propeller. Basic experimental and theoretical results from studies on the unsteady aerodynamics of oscillating cascades are outlined.

  14. Aeroacoustics of advanced propellers

    NASA Technical Reports Server (NTRS)

    Groeneweg, John F.

    1990-01-01

    The aeroacoustics of advanced, high speed propellers (propfans) are reviewed from the perspective of NASA research conducted in support of the Advanced Turboprop Program. Aerodynamic and acoustic components of prediction methods for near and far field noise are summarized for both single and counterrotation propellers in uninstalled and configurations. Experimental results from tests at both takeoff/approach and cruise conditions are reviewed with emphasis on: (1) single and counterrotation model tests in the NASA Lewis 9 by 15 (low speed) and 8 by 6 (high speed) wind tunnels, and (2) full scale flight tests of a 9 ft (2.74 m) diameter single rotation wing mounted tractor and a 11.7 ft (3.57 m) diameter counterrotation aft mounted pusher propeller. Comparisons of model data projected to flight with full scale flight data show good agreement validating the scale model wind tunnel approach. Likewise, comparisons of measured and predicted noise level show excellent agreement for both single and counterrotation propellers. Progress in describing angle of attack and installation effects is also summarized. Finally, the aeroacoustic issues associated with ducted propellers (very high bypass fans) are discussed.

  15. Propellers in Saturn's rings

    NASA Astrophysics Data System (ADS)

    Sremcevic, M.; Stewart, G. R.; Albers, N.; Esposito, L. W.

    2014-04-01

    Theoretical studies and simulations have demonstrated the effects caused by objects embedded in planetary rings [5, 8]. Even if the objects are too small to be directly observed, each creates a much larger gravitational imprint on the surrounding ring material. These strongly depend on the mass of the object and range from "S" like propeller-shaped structures for about 100m-sized icy bodies to the opening of circumferential gaps as in the case of the embedded moons Pan and Daphnis and their corresponding Encke and Keeler Gaps. Since the beginning of the Cassini mission many of these smaller objects (~ 100m in size) have been identified in Saturn's A ring through their propeller signature in the images [10, 7, 9, 11]. Furthermore, recent Cassini observations indicate the possible existence of objects embedded even in Saturn's B and C ring [6, 2]. In this paper we present our new results about by now classical A ring propellers and more enigmatic B ring population. Due to the presence of self-gravity wakes the analysis of propeller brightness in ISS images always bears some ambiguity [7, 9] and consequently the exact morphology of propellers is not a settled issue. In 2008 we obtained a fortunate Cassini Ultraviolet Imaging Spectrograph (UVIS) occultation of the largest A ring propeller Bleriot. Utilizing Cassini ISS images we obtain Bleriot orbit and demonstrate that UVIS Persei Rev42 occultation did cut across Bleriot about 100km downstream from the center. The occultation itself shows a prominent partial gap and higher density outer flanking wakes, while their orientation is consistent with a downstream cut. While in the UVIS occultation the partial gap is more prominent than the flanking wakes, the features mostly seen in Bleriot images are actually flanking wakes. One of the most interesting aspects of the A ring propellers are their wanderings, or longitudinal deviations from a pure circular orbit [11]. We numerically investigated the possibility of simple moon

  16. Propeller noise prediction

    NASA Technical Reports Server (NTRS)

    Zorumski, W. E.

    1983-01-01

    Analytic propeller noise prediction involves a sequence of computations culminating in the application of acoustic equations. The prediction sequence currently used by NASA in its ANOPP (aircraft noise prediction) program is described. The elements of the sequence are called program modules. The first group of modules analyzes the propeller geometry, the aerodynamics, including both potential and boundary layer flow, the propeller performance, and the surface loading distribution. This group of modules is based entirely on aerodynamic strip theory. The next group of modules deals with the actual noise prediction, based on data from the first group. Deterministic predictions of periodic thickness and loading noise are made using Farassat's time-domain methods. Broadband noise is predicted by the semi-empirical Schlinker-Amiet method. Near-field predictions of fuselage surface pressures include the effects of boundary layer refraction and (for a cylinder) scattering. Far-field predictions include atmospheric and ground effects. Experimental data from subsonic and transonic propellers are compared and NASA's future direction is propeller noise technology development are indicated.

  17. Aeroacoustics of advanced propellers

    NASA Astrophysics Data System (ADS)

    Groeneweg, John F.

    The aeroacoustics of advanced, high speed propellers (propfans) are reviewed from the perspective of NASA research conducted in support of the Advanced Turboprop Program. Aerodynamic and acoustic components of prediction methods for near and far field noise are summarized for both single and counterrotation propellers in uninstalled and configurations. Experimental results from tests at both takeoff/approach and cruise conditions are reviewed with emphasis on: (1) single and counterrotation model tests in the NASA Lewis 9 by 15 (low speed) and 8 by 6 (high speed) wind tunnels, and (2) full scale flight tests of a 9 ft (2.74 m) diameter single rotation wing mounted tractor and a 11.7 ft (3.57 m) diameter counterrotation aft mounted pusher propeller. Comparisons of model data projected to flight with full scale flight data show good agreement validating the scale model wind tunnel approach. Likewise, comparisons of measured and predicted noise level show excellent agreement for both single and counterrotation propellers. Progress in describing angle of attack and installation effects is also summarized. Finally, the aeroacoustic issues associated with ducted propellers (very high bypass fans) are discussed.

  18. Solid Propellant Test Article (SPTA) Test Stand

    NASA Technical Reports Server (NTRS)

    1991-01-01

    This photograph shows the Solid Propellant Test Article (SPTA) test stand with the Modified Nasa Motor (M-NASA) test article at the Marshall Space Flight Center (MSFC). The SPTA test stand, 12-feet wide by 12-feet long by 24-feet high, was built in 1989 to provide comparative performance data on nozzle and case insulation material and to verify thermostructural analysis models. A modified NASA 48-inch solid motor (M-NASA motor) with a 12-foot blast tube and 10-inch throat makes up the SPTA. The M-NASA motor is being used to evaluate solid rocket motor internal non-asbestos insulation materials, nozzle designs, materials, and new inspection techniques. New internal motor case instrumentation techniques are also being evaluated.

  19. Theoretical and experimental analysis of the physics of water rockets

    NASA Astrophysics Data System (ADS)

    Barrio-Perotti, R.; Blanco-Marigorta, E.; Fernández-Francos, J.; Galdo-Vega, M.

    2010-09-01

    A simple rocket can be made using a plastic bottle filled with a volume of water and pressurized air. When opened, the air pressure pushes the water out of the bottle. This causes an increase in the bottle momentum so that it can be propelled to fairly long distances or heights. Water rockets are widely used as an educational activity, and several mathematical models have been proposed to investigate and predict their physics. However, the real equations that describe the physics of the rockets are so complicated that certain assumptions are usually made to obtain models that are easier to use. These models provide relatively good predictions but fail in describing the complex physics of the flow. This paper presents a detailed theoretical analysis of the physics of water rockets that concludes with the proposal of a physical model. The validity of the model is checked by a series of field tests. The tests showed maximum differences with predictions of about 6%. The proposed model is finally used to investigate the temporal evolution of some significant variables during the propulsion and flight of the rocket. The experience and procedure described in this paper can be proposed to graduate students and also at undergraduate level if certain simplifications are assumed in the general equations.

  20. 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.

  1. Independent Review of the Failure Modes of F-1 Engine and Propellants System

    NASA Technical Reports Server (NTRS)

    Ray, Paul

    2003-01-01

    The F-1 is the powerful engine, that hurdled the Saturn V launch vehicle from the Earth to the moon on July 16,1969. The force that lifted the rocket overcoming the gravitational force during the first stage of the flight was provided by a cluster of five F-1 rocket engines, each of them developing over 1.5 million pounds of thrust (MSFC-MAN-507). The F-1 Rocket engine used RP-1 (Rocket Propellant-1, commercially known as Kerosene), as fuel with lox (liquid Oxygen) as oxidizer. NASA terminated Saturn V activity and has focused on Space Shuttle since 1972. The interest in rocket system has been revived to meet the National Launch System (NLS) program and a directive from the President to return to the Moon and exploration of the space including Mars. The new program Space Launch Initiative (SLI) is directed to drastically reduce the cost of flight for payloads, and adopt a reusable launch vehicle (RLV). To achieve this goal it is essential to have the ability of lifting huge payloads into low earth orbit. Probably requiring powerful boosters as strap-ons to a core vehicle, as was done for the Saturn launch vehicle. The logic in favor of adopting Saturn system, a proven technology, to meet the SLI challenge is very strong. The F-1 engine was the largest and most powerful liquid rocket engine ever built, and had exceptional performance. This study reviews the failure modes of the F-1 engine and propellant system.

  2. Low-Cost Propellant Launch to Earth Orbit from a Tethered Balloon

    NASA Technical Reports Server (NTRS)

    Wilcox, Brian H.

    2006-01-01

    Propellant will be more than 85% of the mass that needs to be lofted into Low Earth Orbit (LEO) in the planned program of Exploration of the Moon, Mars, and beyond. This paper describes a possible means for launching thousands of tons of propellant per year into LEO at a cost 15 to 30 times less than the current launch cost per kilogram. The basic idea is to mass-produce very simple, small and relatively low-performance rockets at a cost per kilogram comparable to automobiles, instead of the 25X greater cost that is customary for current launch vehicles that are produced in small quantities and which are manufactured with performance near the limits of what is possible. These small, simple rockets can reach orbit because they are launched above 95% of the atmosphere, where the drag losses even on a small rocket are acceptable, and because they can be launched nearly horizontally with very simple guidance based primarily on spin-stabilization. Launching above most of the atmosphere is accomplished by winching the rocket up a tether to a balloon. A fuel depot in equatorial orbit passes over the launch site on every orbit (approximately every 90 minutes). One or more rockets can be launched each time the fuel depot passes overhead, so the launch rate can be any multiple of 6000 small rockets per year, a number that is sufficient to reap the benefits of mass production.

  3. Propellers in yaw

    NASA Technical Reports Server (NTRS)

    Ribner, Herbert S

    1945-01-01

    It was realized as early as 1909 that a propeller in yaw develops a side force like that of a fin. In 1917, R. G. Harris expressed this force in terms of the torque coefficient for the unyawed propeller. Of several attempts to express the side force directly in terms of the shape of the blades, however, none has been completely satisfactory. An analysis that incorporates induction effects not adequately covered in previous work and that gives good agreement with experiment over a wide range of operating conditions is presented. The present analysis shows that the fin analogy may be extended to the form of the side-force expression and that the effective fin area may be taken as the projected side area of the propeller.

  4. 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)

  5. A Plasma Rocket Demonstration on the International Space Station

    NASA Astrophysics Data System (ADS)

    Petro, A.

    2002-01-01

    in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One feature of this concept is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. For this reason the system is called the Variable Specific Impulse Magneto-plasma Rocket or VASIMR. This ability to vary specific impulse and thrust will allow for optimum low thrust interplanetary trajectories and results in shorter trip times than is possible with fixed specific impulse systems while preserving adequate payload margins. demonstrations are envisioned. A ground-based experiment of a low-power VASIMR prototype rocket is currently underway at the Advanced Space Propulsion Laboratory. The next step is a proposal to build and fly a 25-kilowatt VASIMR rocket as an external payload on the International Space Station. This experiment will provide an opportunity to demonstrate the performance of the rocket in space and measure the induced environment. The experiment will also utilize the space station for its intended purpose as a laboratory with vacuum conditions that cannot be matched by any laboratory on Earth. propulsion on the space station. An electric propulsion system like VASIMR, if provided with sufficient electrical power, could provide continuous drag force compensation for the space station. Drag compensation would eliminate the need for reboosting the station, an operation that will consume about 60 metric tons of propellant in a ten-year period. In contrast, an electric propulsion system would require very little propellant. In fact, a system like VASIMR can use waste hydrogen from the station's life support system as its propellant. This waste hydrogen is otherwise dumped overboard. Continuous drag compensation would also improve the microgravity conditions on the station. So electric propulsion can reduce propellant

  6. Optimum propeller wind turbines

    NASA Astrophysics Data System (ADS)

    Sanderson, R. J.; Archer, R. D.

    1983-12-01

    The Prandtl-Betz-Theodorsen theory of heavily loaded airscrews has been adapted to the design of propeller windmills which are to be optimized for maximum power coefficient. It is shown that the simpler, light-loading, constant-area wake assumption can generate significantly different 'optimum' performance and geometry, and that it is therefore not appropriate to the design of propeller wind turbines when operating in their normal range of high-tip-speed-to-wind-speed ratio. Design curves for optimum power coefficient are presented and an example of the design of a typical two-blade optimum rotor is given.

  7. Cryogenic Propellant Scavenging

    NASA Technical Reports Server (NTRS)

    Louie, B.; Kemp, N. J.; Daney, D. E.

    1985-01-01

    A detailed description of a computer model that has been developed for assessing the feasibility of low g cryogen propellant scavenging from the space shuttle External Tank (ET) is given. Either pump-assisted or pressure-induced propellant transfer may be selected. The program will accept a wide range of input variables, including the fuel to be transferred (LOX or LH2), heat leaks, tank temperatures, and piping and equipment specifications. The model has been parametrically analyzed to determine initial design specification for the system.

  8. Origin of how steam rockets can reduce space transport cost by orders of magnitude

    SciTech Connect

    Zuppero, A.; Larson, T.K.; Schnitzler, B.G.; Rice, J.W.; Hill, T.J.; Richins, W.D.; Parlier, L.; Werner, J.E.

    1999-01-01

    A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10{percent} of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon. {copyright} {ital 1999 American Institute of Physics.}

  9. Active synchrophasing of propeller unbalance

    NASA Technical Reports Server (NTRS)

    Kaptein, Dick

    1992-01-01

    The results of a survey are presented to reduce the inflight propeller unbalance vibrations in the cabin of the Fokker 50 airplanes. Several approaches have been investigated. Active synchrophasing of the unbalance vibrations of both propellers appears to be successful.

  10. Notes on propeller design III

    NASA Technical Reports Server (NTRS)

    Munk, Max M

    1922-01-01

    The air flow and the air force created by all elements of the propeller blades lying in a ring located between two concentric circles around the propeller axis are independent of what happens in other rings.

  11. 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

  12. Liquid Methane/Liquid Oxygen Propellant Conditioning Feed System (PCFS) Test Rigs

    NASA Technical Reports Server (NTRS)

    Skaff, A.; Grasl, S.; Nguyen, C.; Hockenberry S.; Schubert, J.; Arrington, L.; Vasek, T.

    2008-01-01

    As part of their Propulsion and Cryogenic Advanced Development (PCAD) program, NASA has embarked upon an effort to develop chemical rocket engines which utilize non-toxic, cryogenic propellants such as liquid oxygen (LO2) and liquid methane (LCH4). This effort includes the development and testing of a 100 lbf Reaction Control Engine (RCE) that will be used to evaluate the performance of a LO2/LCH4 rocket engine over a broad range of propellant temperatures and pressures. This testing will take place at NASA-Glenn Research Center's (GRC) Research Combustion Laboratory (RCL) test facility in Cleveland, OH, and is currently scheduled to begin in late 2008. While the initial tests will be performed at sea level, follow-on testing will be performed at NASA-GRC's Altitude Combustion Stand (ACS) for altitude testing. In support of these tests, Sierra Lobo, Inc. (SLI) has designed, developed, and fabricated two separate portable propellant feed systems under the Propellant Conditioning and Feed System (PCFS) task: one system for LCH4, and one for LO2. These systems will be capable of supplying propellants over a large range of conditions from highly densified to several hundred pounds per square inch (psi) saturated. This paper presents the details of the PCFS design and explores the full capability of these propellant feed systems.

  13. Improved method of measuring pressure coupled response for composite solid propellants

    NASA Astrophysics Data System (ADS)

    Su, Wanxing; Wang, Ningfei; Li, Junwei; Zhao, Yandong; Yan, Mi

    2014-04-01

    Pressure coupled response is one of the main causes of combustion instability in the solid rocket motor. It is also a characteristic parameter for predicting the stability. The pressure coupled response function is usually measured by different methods to evaluate the performance of new propellant. Based on T-burner and "burning surface doubled and secondary attenuation", an improved method for measuring the pressure coupled response of composite propellant is introduced in this article. A computational fluid dynamics (CFD) study has also been conducted to validate the method and to understand the pressure oscillation phenomenon in T-burner. Three rounds of tests were carried out on the same batch of aluminized AP/HTPB composite solid propellant. The experimental results show that the sample propellant had a high response function under the conditions of high pressure (~11.5 MPa) and low frequency (~140 Hz). The numerically predicted oscillation frequency and amplitude are consistent with the experimental results. One practical solid rocket motor using this sample propellant was found to experience pressure oscillation at the end of burning. This confirms that the sample propellant is prone to combustion instability. Finally, acoustic pressure distribution and phase difference in T-burner were analyzed. Both the experimental and numerical results are found to be associated with similar acoustic pressure distribution. And the phase difference analysis showed that the pressure oscillations at the head end of the T-burner are 180° out of phase from those in the aft end of the T-burner.

  14. Long-term firing tests of the nozzles of rocket engines made on the basis of carbon composite materials

    NASA Astrophysics Data System (ADS)

    Gubertov, A. M.; Koshlakov, V. V.; Mironov, V. V.; Rubinskii, V. R.; Pashutov, A. V.; Antipov, Ye. A.; Bratukhin, N. A.; Volkov, N. N.; Volkova, L. I.; Tsatsuev, S. M.; Tlevtsezhev, V. V.

    2012-12-01

    The results of the experimental investigation of the physico-chemical processes of interaction, destruction, and ablation of carbon composite materials and oxidation-protective coatings of nozzles of liquid-propellant rocket engines are presented. The thermally-stressed state of the joint between the nozzle made of composite material and the metallic combustion chamber of the rocket engine under standard operating conditions have been analyzed.

  15. 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

  16. 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.

  17. Small centrifugal pumps for low thrust rockets

    NASA Technical Reports Server (NTRS)

    Gulbrandsen, N. C.; Furst, R. B.; Burgess, R. M.; Scheer, D. D.

    1985-01-01

    This paper presents the results of a combined analytical and experimental investigation of low specific speed pumps for potential use as components of propellant feed systems for low thrust rocket engines. Shrouded impellers and open face impellers were tested in volute type and vaned diffuser type pumps. Full- and partial-emission diffusers and full- and partial-admission impellers were tested. Axial and radial loads, head and efficiency versus flow, and cavitation tests were conducted. Predicted performance of two pumps are compared when pumping water and liquid hydrogen. Detailed pressure loss and parasitic power values are presented for two pump configurations. Partial-emission diffusers were found to permit use of larger impeller and diffuser passages with a minimal performance penalty. Normal manufacturing tolerances were found to result in substantial power requirement variation with only a small pressure rise change. Impeller wear ring leakage was found to reduce pump pressure rise to an increasing degree as the pump flowrate was decreased.

  18. Shuttle rocket booster computational fluid dynamics

    NASA Technical Reports Server (NTRS)

    Chung, T. J.; Park, O. Y.

    1988-01-01

    Additional results and a revised and improved computer program listing from the shuttle rocket booster computational fluid dynamics formulations are presented. Numerical calculations for the flame zone of solid propellants are carried out using the Galerkin finite elements, with perturbations expanded to the zeroth, first, and second orders. The results indicate that amplification of oscillatory motions does indeed prevail in high frequency regions. For the second order system, the trend is similar to the first order system for low frequencies, but instabilities may appear at frequencies lower than those of the first order system. The most significant effect of the second order system is that the admittance is extremely oscillatory between moderately high frequency ranges.

  19. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Shoji, J. M.

    1981-01-01

    An analytical study evaluating thrust chamber cooling engine cycles and preliminary engine design for low thrust chemical rocket engines for orbit transfer vehicles is described. Oxygen/hydrogen, oxygen/methane, and oxygen/RP-1 engines with thrust levels from 444.8 N to 13345 N, and chamber pressures from 13.8 N/sq cm to 689.5 N/sq cm were evaluated. The physical and thermodynamic properties of the propellant theoretical performance data, and transport properties are documented. The thrust chamber cooling limits for regenerative/radiation and film/radiation cooling are defined and parametric heat transfer data presented. A conceptual evaluation of a number of engine cycles was performed and a 2224.1 N oxygen/hydrogen engine cycle configuration and a 2224.1 N oxygen/methane configuration chosen for preliminary engine design. Updated parametric engine data, engine design drawings, and an assessment of technology required are presented.

  20. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    NASA Technical Reports Server (NTRS)

    Jones, Carl P.; Robertson, Elizabeth H.; Koelbl, Mary Beth; Singer, Chris

    2016-01-01

    Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.