Research on shock wave characteristics in the isolator of central strut rocket-based combined cycle engine under Ma5.5

NASA Astrophysics Data System (ADS)

Wei, Xianggeng; Xue, Rui; Qin, Fei; Hu, Chunbo; He, Guoqiang

2017-11-01

A numerical calculation of shock wave characteristics in the isolator of central strut rocket-based combined cycle (RBCC) engine fueled by kerosene was carried out in this paper. A 3D numerical model was established by the DES method. The kerosene chemical kinetic model used the 9-component and 12-step simplified mechanism model. Effects of fuel equivalence ratio, inflow total temperature and central strut rocket on-off on shock wave characteristics were studied under Ma5.5. Results demonstrated that with the increase of equivalence ratio, the leading shock wave moves toward upstream, accompanied with higher possibility of the inlet unstart. However, the leading shock wave moves toward downstream as the inflow total temperature rises. After the central strut rocket is closed, the leading shock wave moves toward downstream, which can reduce risks of the inlet unstart. State of the shear layer formed by the strut rocket jet flow and inflow can influence the shock train structure significantly.

Direct launch using the electric rail gun

NASA Technical Reports Server (NTRS)

Barber, J. P.

1983-01-01

The concept explored involves using a large single stage electric rail gun to achieve orbital velocities. Exit aerodynamics, launch package design and size, interior ballistics, system and component sizing and design concepts are treated. Technology development status and development requirements are identified and described. The expense of placing payloads in Earth orbit using conventional chemical rockets is considerable. Chemical rockets are very inefficient in converting chemical energy into payload kinetic energy. A rocket motor is relatively expensive and is usually expended on each launch. In addition specialized and expensive forms of fuel are required. Gun launching payloads directly to orbit from the Earth's surface is a possible alternative. Guns are much more energy efficient than rockets. The high capital cost of the gun installation can be recovered by reusing it over and over again. Finally, relatively inexpensive fuel and large quantities of energy are readily available to a fixed installation on the Earth's surface.

Test Results of the RS-44 Integrated Component Evaluator Liquid Oxygen/Hydrogen Rocket Engine

NASA Technical Reports Server (NTRS)

Sutton, R. F.; Lariviere, B. W.

1993-01-01

An advanced LOX/LH2 expander cycle rocket engine, producing 15,000 lbf thrust for Orbital Transfer Vehicle missions, was tested to determine ignition, transition, and main stage characteristics. Detail design and fabrication of the pump fed RS44 integrated component evaluator (ICE) was accomplished using company discretionary resources and was tested under this contracted effort. Successful demonstrations were completed to about the 50 percent fuel turbopump power level (87,000 RPM), but during this last test, a high pressure fuel turbopump (HPFTP) bearing failed curtailing the test program. No other hardware were affected by the HPFTP premature shutdown. The ICE operations matched well with the predicted start transient simulations. The tests demonstrated the feasibility of a high performance advanced expander cycle engine. All engine components operated nominally, except for the HPFTP, during the engine hot-fire tests. A failure investigation was completed using company discretionary resources.

Design and testing of digitally manufactured paraffin Acrylonitrile-butadiene-styrene hybrid rocket motors

NASA Astrophysics Data System (ADS)

McCulley, Jonathan M.

This research investigates the application of additive manufacturing techniques for fabricating hybrid rocket fuel grains composed of porous Acrylonitrile-butadiene-styrene impregnated with paraffin wax. The digitally manufactured ABS substrate provides mechanical support for the paraffin fuel material and serves as an additional fuel component. The embedded paraffin provides an enhanced fuel regression rate while having no detrimental effect on the thermodynamic burn properties of the fuel grain. Multiple fuel grains with various ABS-to-Paraffin mass ratios were fabricated and burned with nitrous oxide. Analytical predictions for end-to-end motor performance and fuel regression are compared against static test results. Baseline fuel grain regression calculations use an enthalpy balance energy analysis with the material and thermodynamic properties based on the mean paraffin/ABS mass fractions within the fuel grain. In support of these analytical comparisons, a novel method for propagating the fuel port burn surface was developed. In this modeling approach the fuel cross section grid is modeled as an image with white pixels representing the fuel and black pixels representing empty or burned grid cells.

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 ceramic composite of mixed hafnium carbide and tantalum carbide reinforced with graphite fibers.

B-1 and B-3 Test Stands at NASA’s Plum Brook Station

NASA Image and Video Library

1966-09-21

Operation of the High Energy Rocket Engine Research Facility (B-1), left, and Nuclear Rocket Dynamics and Control Facility (B-3) at the National Aeronautics and Space Administration’s (NASA) Plum Brook Station in Sandusky, Ohio. The test stands were constructed in the early 1960s to test full-scale liquid hydrogen fuel systems in simulated altitude conditions. Over the next decade each stand was used for two major series of liquid hydrogen rocket tests: the Nuclear Engine for Rocket Vehicle Application (NERVA) and the Centaur second-stage rocket program. The different components of these rocket engines could be studied under flight conditions and adjusted without having to fire the engine. Once the preliminary studies were complete, the entire engine could be fired in larger facilities. The test stands were vertical towers with cryogenic fuel and steam ejector systems. B-1 was 135 feet tall, and B-3 was 210 feet tall. Each test stand had several levels, a test section, and ground floor shop areas. The test stands relied on an array of support buildings to conduct their tests, including a control building, steam exhaust system, and fuel storage and pumping facilities. A large steam-powered altitude exhaust system reduced the pressure at the exhaust nozzle exit of each test stand. This allowed B-1 and B-3 to test turbopump performance in conditions that matched the altitudes of space.

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

NASA Astrophysics Data System (ADS)

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

2016-04-01

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

Nuclear Thermal Rocket Element Environmental Simulator (NTREES) Phase II Upgrade Activities

NASA Technical Reports Server (NTRS)

Emrich, William J.; Moran, Robert P.; Pearson, J. Bose

2013-01-01

To support the on-going nuclear thermal propulsion effort, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The facility to perform this testing is referred to as the Nuclear Thermal Rocket Element Environment Simulator (NTREES). This device can simulate the environmental conditions (minus the radiation) to which nuclear rocket fuel components will be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner so as to accurately reproduce the temperatures and heat fluxes which would normally occur as a result of nuclear fission and would be exposed to flowing hydrogen. Initial testing of a somewhat prototypical fuel element has been successfully performed in NTREES and the facility has now been shutdown to allow for an extensive reconfiguration of the facility which will result in a significant upgrade in its capabilities. Keywords: Nuclear Thermal Propulsion, Simulator

Air Force Research Laboratory Technology Milestones 2007

DTIC Science & Technology

2007-01-01

Propulsion Fuel Pumps and Fuel Systems Liquid Rockets and Combustion Gas Generators Micropropulsion Gears Monopropellants High-Cycle Fatigue and Its... Systems Electric Propulsion Engine Health Monitoring Systems High-Energy-Density Matter Exhaust Nozzles Injectors and Spray Measurements Fans Laser...of software models to drive development of component-based systems and lightweight domain-specific specification and verification technology. Highly

Fluidized-Solid-Fuel Injection Process

NASA Technical Reports Server (NTRS)

Taylor, William

1992-01-01

Report proposes development of rocket engines burning small grains of solid fuel entrained in gas streams. Main technical discussion in report divided into three parts: established fluidization technology; variety of rockets and rocket engines used by nations around the world; and rocket-engine equation. Discusses significance of specific impulse and ratio between initial and final masses of rocket. Concludes by stating three important reasons to proceed with new development: proposed engines safer; fluidized-solid-fuel injection process increases variety of solid-fuel formulations used; and development of fluidized-solid-fuel injection process provides base of engineering knowledge.

Supercomputer modeling of hydrogen combustion in rocket engines

NASA Astrophysics Data System (ADS)

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

2013-08-01

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

Kerosene-Fuel Engine Testing Under Way

NASA Image and Video Library

2003-11-17

NASA Stennis Space Center engineers conducted a successful cold-flow test of an RS-84 engine component Sept. 24. The RS-84 is a reusable engine fueled by rocket propellant - a special blend of kerosene - designed to power future flight vehicles. Liquid oxygen was blown through the RS-84 subscale preburner to characterize the test facility's performance and the hardware's resistance. Engineers are now moving into the next phase, hot-fire testing, which is expected to continue into February 2004. The RS-84 engine prototype, developed by the Rocketdyne Propulsion and Power division of The Boeing Co. of Canoga Park, Calif., is one of two competing Rocket Engine Prototype technologies - a key element of NASA's Next Generation Launch Technology program.

Kerosene-Fuel Engine Testing Under Way

NASA Technical Reports Server (NTRS)

2003-01-01

NASA Stennis Space Center engineers conducted a successful cold-flow test of an RS-84 engine component Sept. 24. The RS-84 is a reusable engine fueled by rocket propellant - a special blend of kerosene - designed to power future flight vehicles. Liquid oxygen was blown through the RS-84 subscale preburner to characterize the test facility's performance and the hardware's resistance. Engineers are now moving into the next phase, hot-fire testing, which is expected to continue into February 2004. The RS-84 engine prototype, developed by the Rocketdyne Propulsion and Power division of The Boeing Co. of Canoga Park, Calif., is one of two competing Rocket Engine Prototype technologies - a key element of NASA's Next Generation Launch Technology program.

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

NASA Technical Reports Server (NTRS)

Jones, Ronald E.; Carter, Robert W.

2012-01-01

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

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

NASA Technical Reports Server (NTRS)

Schoenman, L.

1981-01-01

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

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

NASA Technical Reports Server (NTRS)

Jaffe, Richard L.

2014-01-01

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

Computational Flow Analysis of a Left Ventricular Assist Device

NASA Technical Reports Server (NTRS)

Kiris, Cetin; Kwak, Dochan; Benkowski, Robert

1995-01-01

Computational fluid dynamics has been developed to a level where it has become an Indispensable part of aerospace research and design. Technology developed foe aerospace applications am also be utilized for the benefit of human health. For example, a flange-to-flange rocket engine fuel-pump simulation includes the rotating and non-rotating components: the flow straighteners, the impeller, and diffusers A Ventricular Assist Device developed by NASA Johnson Space Center and Baylor College of Medicine has a design similar to a rocket engine fuel pump in that it also consists of a flow straightener, an impeller, and a diffuser. Accurate and detailed knowledge of the flowfield obtained by incompressible flow calculations can be greatly beneficial to designers in their effort to reduce the cost and improve the reliability of these devices. In addition to the geometric complexities, a variety of flow phenomena are encountered in biofluids Then include turbulent boundary layer separation, wakes, transition, tip vortex resolution, three-dimensional effects, and Reynolds number effects. In order to increase the role of Computational Fluid Dynamics (CFD) in the design process the CFD analysis tools must be evaluated and validated so that designers gain Confidence in their use. The incompressible flow solver, INS3D, has been applied to flow inside of a liquid rocket engine turbopump components and extensively validated. This paper details how the computational flow simulation capability developed for liquid rocket engine pump component analysis has bean applied to the Left Ventricular Assist Device being developed jointly by NASA JSC and Baylor College of Medicine.

Engineers demonstrate the pocket rocket

NASA Technical Reports Server (NTRS)

1996-01-01

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

Response Surface Modeling of Combined-Cycle Propulsion Components using Computational Fluid Dynamics

NASA Technical Reports Server (NTRS)

Steffen, C. J., Jr.

2002-01-01

Three examples of response surface modeling with CFD are presented for combined cycle propulsion components. The examples include a mixed-compression-inlet during hypersonic flight, a hydrogen-fueled scramjet combustor during hypersonic flight, and a ducted-rocket nozzle during all-rocket flight. Three different experimental strategies were examined, including full factorial, fractionated central-composite, and D-optimal with embedded Plackett-Burman designs. The response variables have been confined to integral data extracted from multidimensional CFD results. Careful attention to uncertainty assessment and modeling bias has been addressed. The importance of automating experimental setup and effectively communicating statistical results are emphasized.

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

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

Nitrous Oxide/Paraffin Hybrid Rocket Engines

NASA Technical Reports Server (NTRS)

Zubrin, Robert; Snyder, Gary

2010-01-01

Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

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.

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

NASA Technical Reports Server (NTRS)

Miller, Riley O

1950-01-01

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

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.

Hydrocarbon Fuel Thermal Performance Modeling based on Systematic Measurement and Comprehensive Chromatographic Analysis

DTIC Science & Technology

2016-07-27

is a common requirement for aircraft, rockets , and hypersonic vehicles. The Aerospace Fuels Quality Test and Model Development (AFQTMoDev) project...was initiated to mature fuel quality assurance practices for rocket grade kerosene, thereby ensuring operational readiness of conventional and...and reliability, is a common requirement for aircraft, rockets , and hypersonic vehicles. The Aerospace Fuels Quality Test and Model Development

SLS Resource Reel Aug 2016 orig

NASA Image and Video Library

2016-07-04

Space Launch System Resource Reel Description: This video includes launch animation of NASA’s Space Launch System (SLS), as well as work taking place across NASA centers and the country to build and test the various components that make up the rocket including: the 5-segment solid rocket boosters, the RS-25 rocket engines, the massive tanks that make up the Core Stage of the rocket that fuels the RS-25 engines, and upper portions of the rocket that connect the interim cryogenic propulsion stage to the Orion spacecraft. SLS, is an advanced launch vehicle for a new era of exploration beyond Earth’s orbit into deep space. SLS, the world’s most powerful rocket, will launch astronauts in the agency’s Orion spacecraft on missions to an asteroid and eventually to Mars, while opening new possibilities for other payloads including robotic scientific missions to places like Mars, Saturn and Jupiter. Graphic Information: PAO Name:Kim Henry Phone Number:256-544-1899 Email Address: kimberly.m.henry@nasa.gov

Design considerations for a pressure-driven multi-stage rocket

NASA Astrophysics Data System (ADS)

Sauerwein, Steven Craig

2002-01-01

The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior of the pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large.

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.

Space Shuttle Projects

NASA Image and Video Library

2001-01-01

The Space Shuttle represented an entirely new generation of space vehicles, the world's first reusable spacecraft. Unlike earlier expendable rockets, the Shuttle was designed to be launched over and over again and would serve as a system for ferrying payloads and persornel to and from Earth orbit. The Shuttle's major components are the orbiter spacecraft; the three main engines, with a combined thrust of more than 1.2 million pounds; the huge external tank (ET) that feeds the liquid hydrogen fuel and liquid oxygen oxidizer to the three main engines; and the two solid rocket boosters (SRB's), with their combined thrust of some 5.8 million pounds, that provide most of the power for the first two minutes of flight. Crucially involved with the Space Shuttle program virtually from its inception, the Marshall Space Flight Center (MSFC) played a leading role in the design, development, testing, and fabrication of many major Shuttle propulsion components. The MSFC was assigned responsibility for developing the Shuttle orbiter's high-performance main engines, the most complex rocket engines ever built. The MSFC was also responsible for developing the Shuttle's massive ET and the solid rocket motors and boosters.

Space Shuttle Projects

NASA Image and Video Library

1975-01-01

The Space Shuttle represented an entirely new generation of space vehicle, the world's first reusable spacecraft. Unlike earlier expendable rockets, the Shuttle was designed to be launched over and over again and would serve as a system for ferrying payloads and persornel to and from Earth orbit. The Shuttle's major components are the orbiter spacecraft; the three main engines, with a combined thrust of more than 1.2 million pounds; the huge external tank (ET) that feeds the liquid hydrogen fuel and liquid oxygen oxidizer to the three main engines; and the two solid rocket boosters (SRB's), with their combined thrust of some 5.8 million pounds. The SRB's provide most of the power for the first two minutes of flight. Crucially involved with the Space Shuttle program virtually from its inception, the Marshall Space Flight Center (MSFC) played a leading role in the design, development, testing, and fabrication of many major Shuttle propulsion components. The MSFC was assigned responsibility for developing the Shuttle orbiter's high-performance main engines, the most complex rocket engines ever built. The MSFC was also responsible for developing the Shuttle's massive ET and the solid rocket motors and boosters.

Hydrocarbon Fuel Thermal Performance Modeling based on Systematic Measurement and Comprehensive Chromatographic Analysis

DTIC Science & Technology

2016-07-31

fueled liquid rocket engine, enthalpy is removed from the combustion chamber by a regenerative cooling system comprising a series of passages through... rocket engine, enthalpy is removed from the combustion chamber by a regenerative cooling system comprising a series of passages through which fuel flows...the unprecedented correlation of comprehensive two-dimensional gas chromatographic (GC×GC) rocket fuel data with physical and thermochemical

REMOTE SENSING OF PERCHLORATE EFFECTS ON SALT CEDAR PRELIMINARY RESULTS FROM THE LAS VEGAS WASH

EPA Science Inventory

Sodium Perchlorate and ammonium Perchlorate, major components of solid rocket fuel, have been manufactured in the Las Vegas Valley immediately up gradient from the Las Vegas Wash, since 1945 and 1956, respectively. Measurements of emerging ground water quality in the vici...

Hydrocarbon-Fueled Rocket Engine Plume Diagnostics: Analytical Developments and Experimental Results

NASA Technical Reports Server (NTRS)

Tejwani, Gopal D.; McVay, Gregory P.; Langford, Lester A.; St. Cyr, William W.

2006-01-01

A viewgraph presentation describing experimental results and analytical developments about plume diagnostics for hydrocarbon-fueled rocket engines is shown. The topics include: 1) SSC Plume Diagnostics Background; 2) Engine Health Monitoring Approach; 3) Rocket Plume Spectroscopy Simulation Code; 4) Spectral Simulation for 10 Atomic Species and for 11 Diatomic Molecular Electronic Bands; 5) "Best" Lines for Plume Diagnostics for Hydrocarbon-Fueled Rocket Engines; 6) Experimental Set Up for the Methane Thruster Test Program and Experimental Results; and 7) Summary and Recommendations.

Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

NASA Astrophysics Data System (ADS)

Emrich, William J.

2008-01-01

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

Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.

2008-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2013-01-01

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

ACCUMULATION AND FATE OF PERCHLORATE IN PLANTS

EPA Science Inventory

Perchlorate, a component of solid rocket fuels, has emerged as a potential threat to surface water and groundwater at several locations in the U.S. Perchlorate levels up to 16 ug/L were detected in Lake Mead and 5-9 ug/L in the lower Colorado River. The water from the Colorado Ri...

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.

Around Marshall

NASA Image and Video Library

1998-11-04

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.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Youngblood, Stewart

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 ofmore » 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.« less

Response of selected plant and insect species to simulated SRM exhaust mixtures and to exhaust components from SRM fuels

NASA Technical Reports Server (NTRS)

Heck, W. W.

1980-01-01

The possible biologic effects of exhaust products from solid rocket motor (SRM) burns associated with the space shuttle are examined. The major components of the exhaust that might have an adverse effect on vegetation, HCl and Al2O3 are studied. Dose response curves for native and cultivated plants and selected insects exposed to simulated exhaust and component chemicals from SRM exhaust are presented. A system for dispensing and monitoring component chemicals of SRM exhaust (HCl and Al2O3) and a system for exposing test plants to simulated SRM exhaust (controlled fuel burns) are described. The effects of HCl, Al2O3, and mixtures of the two on the honeybee, the corn earworm, and the common lacewing and the effects of simulated exhaust on the honeybee are discussed.

Understanding Kelvin-Helmholtz instability in paraffin-based hybrid rocket fuels

NASA Astrophysics Data System (ADS)

Petrarolo, Anna; Kobald, Mario; Schlechtriem, Stefan

2018-04-01

Liquefying fuels show higher regression rates than the classical polymeric ones. They are able to form, along their burning surface, a low viscosity and surface tension liquid layer, which can become unstable (Kelvin-Helmholtz instability) due to the high velocity gas flow in the fuel port. This causes entrainment of liquid droplets from the fuel surface into the oxidizer gas flow. To better understand the droplets entrainment mechanism, optical investigations on the combustion behaviour of paraffin-based hybrid rocket fuels in combination with gaseous oxygen have been conducted in the framework of this research. Combustion tests were performed in a 2D single-slab burner at atmospheric conditions. High speed videos were recorded and analysed with two decomposition techniques. Proper orthogonal decomposition (POD) and independent component analysis (ICA) were applied to the scalar field of the flame luminosity. The most excited frequencies and wavelengths of the wave-like structures characterizing the liquid melt layer were computed. The fuel slab viscosity and the oxidizer mass flow were varied to study their influence on the liquid layer instability process. The combustion is dominated by periodic, wave-like structures for all the analysed fuels. Frequencies and wavelengths characterizing the liquid melt layer depend on the fuel viscosity and oxidizer mass flow. Moreover, for very low mass flows, no wavelength peaks are detected for the higher viscosity fuels. This is important to better understand and predict the onset and development of the entrainment process, which is connected to the amplification of the longitudinal waves.

Heat Transfer and Thermal Stability Research for Advanced Hydrocarbon Fuel Technologies

NASA Technical Reports Server (NTRS)

DeWitt, Kenneth; Stiegemeier, Benjamin

2005-01-01

In recent years there has been increased interest in the development of a new generation of high performance boost rocket engines. These efforts, which will represent a substantial advancement in boost engine technology over that developed for the Space Shuttle Main Engines in the early 1970s, are being pursued both at NASA and the United States Air Force. NASA, under its Space Launch Initiative s Next Generation Launch Technology Program, is investigating the feasibility of developing a highly reliable, long-life, liquid oxygen/kerosene (RP-1) rocket engine for launch vehicles. One of the top technical risks to any engine program employing hydrocarbon fuels is the potential for fuel thermal stability and material compatibility problems to occur under the high-pressure, high-temperature conditions required for regenerative fuel cooling of the engine combustion chamber and nozzle. Decreased heat transfer due to carbon deposits forming on wetted fuel components, corrosion of materials common in engine construction (copper based alloys), and corrosion induced pressure drop increases have all been observed in laboratory tests simulating rocket engine cooling channels. To mitigate these risks, the knowledge of how these fuels behave in high temperature environments must be obtained. Currently, due to the complexity of the physical and chemical process occurring, the only way to accomplish this is empirically. Heated tube testing is a well-established method of experimentally determining the thermal stability and heat transfer characteristics of hydrocarbon fuels. The popularity of this method stems from the low cost incurred in testing when compared to hot fire engine tests, the ability to have greater control over experimental conditions, and the accessibility of the test section, facilitating easy instrumentation. These benefits make heated tube testing the best alternative to hot fire engine testing for thermal stability and heat transfer research. This investigation used the Heated Tube Facility at the NASA Glenn Research Center to perform a thermal stability and heat transfer characterization of RP-1 in an environment simulating that of a high chamber pressure, regenerative cooled rocket engine. The first step in the research was to investigate the carbon deposition process of previous heated tube experiments by performing scanning electron microscopic analysis in conjunction with energy dispersive spectroscopy on the tube sections. This analysis gave insight into the carbon deposition process and the effect that test conditions played in the formation of deleterious coke. Furthermore, several different formations were observed and noted. One other crucial finding of this investigation was that in sulfur containing hydrocarbon fuels, the interaction of the sulfur components with copper based wall materials presented a significant corrosion problem. This problem in many cases was more life limiting than those posed by the carbon deposition process. The results of this microscopic analysis was detailed and presented at the December 2003 JANNAF Air-Breathing Propulsion Meeting as a Materials Compatibility and Thermal Stability Analysis of common Hydrocarbon Fuels (reference 1).

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.

Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Emrich, William J. Jr.

2008-01-21

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

Predicting the velocity and azimuth of fragments generated by the range destruction or random failure of rocket casings and tankage

NASA Technical Reports Server (NTRS)

Eck, Marshall; Mukunda, Meera

1988-01-01

A calculational method is described which provides a powerful tool for predicting solid rocket motor (SRM) casing and liquid rocket tankage fragmentation response. The approach properly partitions the available impulse to each major system-mass component. It uses the Pisces code developed by Physics International to couple the forces generated by an Eulerian-modeled gas flow field to a Lagrangian-modeled fuel and casing system. The details of the predictive analytical modeling process and the development of normalized relations for momentum partition as a function of SRM burn time and initial geometry are discussed. Methods for applying similar modeling techniques to liquid-tankage-overpressure failures are also discussed. Good agreement between predictions and observations are obtained for five specific events.

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.

The Technique for CFD-Simulation of Fuel Valve from Pneumatic-Hydraulic System of Liquid-Propellant Rocket Engine

NASA Astrophysics Data System (ADS)

Shabliy, L. S.; Malov, D. V.; Bratchinin, D. S.

2018-01-01

In the article the description of technique for simulation of valves for pneumatic-hydraulic system of liquid-propellant rocket engine (LPRE) is given. Technique is based on approach of computational hydrodynamics (Computational Fluid Dynamics - CFD). The simulation of a differential valve used in closed circuit LPRE supply pipes of fuel components is performed to show technique abilities. A schematic and operation algorithm of this valve type is described in detail. Also assumptions made in the construction of the geometric model of the hydraulic path of the valve are described in detail. The calculation procedure for determining valve hydraulic characteristics is given. Based on these calculations certain hydraulic characteristics of the valve are given. Some ways of usage of the described simulation technique for research the static and dynamic characteristics of the elements of the pneumatic-hydraulic system of LPRE are proposed.

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.

Nuclear Propulsion for Space, Understanding the Atom Series.

ERIC Educational Resources Information Center

Corliss, William R.; Schwenk, Francis C.

The operation of nuclear rockets with respect both to rocket theory and to various fuels is described. The development of nuclear reactors for use in nuclear rocket systems is provided, with the Kiwi and NERVA programs highlighted. The theory of fuel element and reactor construction and operation is explained with particular reference to rocket…

76 FR 19768 - Agency Information Collection Activities; Submission to OMB for Review and Approval; Comment...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-04-08

... Activities; Submission to OMB for Review and Approval; Comment Request; NESHAP for Beryllium Rocket Motor... . Title: NESHAP for Beryllium Rocket Motor Fuel Firing (Renewal). ICR Numbers: EPA ICR Number 1125.06, OMB...: Owners or operators of beryllium rocket motor fuel firing facilities. Estimated Number of Respondents: 1...

Bonded foil pressure transducers

NASA Astrophysics Data System (ADS)

Daube, Bernie W.

The design of bonded-foil pressure transducers is discussed, with consideration given to individual components of both the electrical and the mechanical sections of the bonded-foil pressure transducers, as well as to the temperature control and the accuracy specification of these devices. Particular attention is given to applications of bonded foil pressure transducers, which include solid and liquid rocket engine testing for fuel and exhaust pressures, fuel and oil pressure monitoring on jet engines, and nuclear underground safety system pressure monitoring and nuclear test monitoring. A diagram of a transducer cutaway view is included.

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 performance propellants were evaluated: Space storable propellants, including liquid oxygen (LOX) as the oxidizer with nitrogen hydrides or hydrocarbon as fuels. Specifically, a LOX/hydrazine engine was designed, fabricated, and shown to have a 95 pct theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1. Further performance improvment can be obtained by the use of LOX/hydrogen propellants, especially for manned spacecraft applications, and specific designs must be developed and advanced through flight qualification.

Early Rockets

NASA Image and Video Library

1940-03-21

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

Study of advanced techniques for determining the long term performance of components

NASA Technical Reports Server (NTRS)

1973-01-01

The application of existing and new technology to the problem of determining the long-term performance capability of liquid rocket propulsion feed systems is discussed. The long term performance of metal to metal valve seats in a liquid propellant fuel system is stressed. The approaches taken in conducting the analysis are: (1) advancing the technology of characterizing components through the development of new or more sensitive techniques and (2) improving the understanding of the physical of degradation.

Metal hydride and pyrophoric fuel additives for dicyclopentadiene based hybrid propellants

NASA Astrophysics Data System (ADS)

Shark, Steven C.

The purpose of this study is to investigate the use of reactive energetic fuel additives that have the potential to increase the combustion performance of hybrid rocket propellants in terms of solid fuel regression rate and combustion efficiency. Additives that can augment the combustion flame zone in a hybrid rocket motor by means of increased energy feedback to the fuel grain surface are of great interest. Metal hydrides have large volumetric hydrogen densities, which gives these materials high performance potential as fuel additives in terms of specifc impulse. The excess hydrogen and corresponding base metal may also cause an increase in the hybrid rocket solid fuel regression rate. Pyrophoric additives also have potential to increase the solid fuel regression rate by reacting more readily near the burning fuel surface providing rapid energy feedback. An experimental performance evaluation of metal hydride fuel additives for hybrid rocket motor propulsion systems is examined in this study. Hypergolic ignition droplet tests and an accelerated aging study revealed the protection capabilities of Dicyclopentadiene (DCPD) as a fuel binder, and the ability for unaided ignition. Static hybrid rocket motor experiments were conducted using DCPD as the fuel. Sodium borohydride (NabH4) and aluminum hydride (AlH3) were examined as fuel additives. Ninety percent rocket grade hydrogen peroxide (RGHP) was used as the oxidizer. In this study, the sensitivity of solid fuel regression rate and characteristic velocity (C*) efficiency to total fuel grain port mass flux and particle loading is examined. These results were compared to HTPB combustion performance as a baseline. Chamber pressure histories revealed steady motor operation in most tests, with reduced ignition delays when using NabH4 as a fuel additive. The addition of NabH4 and AlH3 produced up to a 47% and 85% increase in regression rate over neat DCPD, respectively. For all test conditions examined C* efficiency ranges between 80% and 90%. The regression rate and C* efficiency mass flux dependence indicate a shift towards a more diffusion controlled system with metal hydride particle addition. Although these types of energetic particles have potential as high performing fuel additives, they can be in low supply and expensive. An opposed flow burner was investigated as a means to screen and characterize hybrid rocket fuels prior to full scale rocket motor testing. Although this type of configuration has been investigated in the past, no comparison has been made to hybrid rocket motor operation in terms of mass flux. Polymeric fuels and low melt temperature fuels with and without additives were investigated via an opposed flow burner. The effects of laminar and turbulent flow regimes on the convective heat transfer in the opposed flow system was depicted in the regression rate trends of these fuels. Regression rate trends similar to hybrid rocket motor operation were depicted, including the entrainment mechanism for paran fuel. However, there was a shift in overall magnitude of these results. A decrease in regression rate occurred for HTPB loaded with passivated nano-aluminum, due to low resonance time in the reaction zone. Previous results have shown that pyrophoric additives can cause an increase in regression rate in the opposed flow burner configuration. It is proposed that the opposed burner is useful as a screening and characterization tool for some propellant combinations. Gaseous oxygen (GOX) was investigated as an oxidizer for similar fuels evaluated with RGHP. Specifically, combustion performance sensitivity to mass flux and MH particle size was investigated. Similar results to the RGHP experiments were observed for the regression rate tends of HTPB, DPCD, and NabH 4 addition. Kinetically limited regression rate dependence on mass flux was observed at the higher mass flux levels. No major increase in C* efficiency was observed for MH addition. The C* efficiency varied with equivalence ratio by approximately 10 percentage points, which was not observed in the RGHP experiments. A 10 percentage point decrease in C* efficiency was observed with increasing mass flux in the system. This was most likely due to poorly mixed fuel and oxidizer in center of the combustion chamber at the higher mass flux levels. Detailed measurements of the hybrid rocket combustion zone is useful for understanding the mechanisms governing performance, but can be difficult to obtain. Traditional slab burner configurations have proven useful but are operationally limited in pressure and mass flux ranges. A new optical cylindrical combustor (OCC) design is presented that allows surface and flame zone imaging and tracking during hybrid rocket motor operation at appreciable mass flux and pressure levels, > 100 kg/s/m2 and > 0.69 MPa. The flame height and regression rate sensitivity to mass flux and chamber pressure was examined for the same fuels examined in the GOX hybrid rocket motor, with the addition of DCPD fuel loaded with Al and unpassivated mechanically activated Al-PTFE. The regression rate trends were on the same order of magnitude of traditional hybrid rocket motor results. A flame height decrease was observed for increased mass flux. The flame height increased with NabH 4 addition, which is most likely a function of increased blowing at the surface. There was no appreciable flame height sensitivity to NabH4 particle size. There was no relative change in flame height or regression rate between the Al and AL-PTFE addition. The OCC allowed visualization of the hybrid rocket fuel flame zone at mass flux and pressure levels that are not known to be report for traditional slab burner configurations in literature. The OCC proved to be a new useful tool for investigated hybrid rocket propellant combustion characteristics.

Dumbo: A pachydermal rocket motor

NASA Technical Reports Server (NTRS)

Kirk, Bill

1991-01-01

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

Dr. Robert H. Goddard and His Rocket

NASA Technical Reports Server (NTRS)

1940-01-01

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

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

PubMed

Hu, Jichao; Chang, Juntao; Bao, Wen

2014-01-01

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

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

PubMed Central

2014-01-01

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

Initial Operation of the Nuclear Thermal Rocket Element Environmental Simulator

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.; Pearson, J. Boise; Schoenfeld, Michael P.

2015-01-01

The Nuclear Thermal Rocket Element Environmental Simulator (NTREES) facility is designed to perform realistic non-nuclear testing of nuclear thermal rocket (NTR) fuel elements and fuel materials. Although the NTREES facility cannot mimic the neutron and gamma environment of an operating NTR, it can simulate the thermal hydraulic environment within an NTR fuel element to provide critical information on material performance and compatibility. The NTREES facility has recently been upgraded such that the power capabilities of the facility have been increased significantly. At its present 1.2 MW power level, more prototypical fuel element temperatures nay now be reached. The new 1.2 MW induction heater consists of three physical units consisting of a transformer, rectifier, and inverter. This multiunit arrangement facilitated increasing the flexibility of the induction heater by more easily allowing variable frequency operation. Frequency ranges between 20 and 60 kHz can accommodated in the new induction heater allowing more representative power distributions to be generated within the test elements. The water cooling system was also upgraded to so as to be capable of removing 100% of the heat generated during testing In this new higher power configuration, NTREES will be capable of testing fuel elements and fuel materials at near-prototypic power densities. As checkout testing progressed and as higher power levels were achieved, several design deficiencies were discovered and fixed. Most of these design deficiencies were related to stray RF energy causing various components to encounter unexpected heating. Copper shielding around these components largely eliminated these problems. Other problems encountered involved unexpected movement in the coil due to electromagnetic forces and electrical arcing between the coil and a dummy test article. The coil movement and arcing which were encountered during the checkout testing effectively destroyed the induction coil in use at the time and resulted in NTREES being out of commission for a couple of months while a new stronger coil was procured. The new coil includes several additional pieces of support structure to prevent coil movement in the future. In addition, new insulating test article support components have been fabricated to prevent unexpected arcing to the test articles. Additional activities are also now underway to address ways in which the radial temperature profiles across test articles may be controlled such that they are more prototypical of what they would encounter in an operating nuclear engine. The causes of the temperature distribution problem are twofold. First, the fuel element test article is isolated in NTREES as opposed to being in the midst of many other mostly identical fuel elements in a nuclear engine. As a result, the fuel element heat flux boundary conditions in NTREES are far from adiabatic as would normally be the case in a reactor. Second, induction heating skews the power distribution such that power is preferentially deposited near the outside of the fuel element. Nuclear heating, conversely, deposits its power much more uniformly throughout the fuel element. Current studies are now looking at various schemes to adjust the amount of thermal radiation emitted from the fuel element surface so as to essentially vary the thermal boundary conditions on the test article. It is hoped that by properly adjusting the thermal boundary conditions on the fuel element test article, it may be possible to substantially correct for the inappropriate radial power distributions resulting from the induction heating so as to yield a more nearly correct temperature distribution throughout the fuel element.

High regression rate hybrid rocket fuel grains with helical port structures

NASA Astrophysics Data System (ADS)

Walker, Sean D.

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

Computation of incompressible viscous flows through turbopump components

NASA Technical Reports Server (NTRS)

Kiris, Cetin; Chang, Leon

1993-01-01

Flow through pump components, such as an inducer and an impeller, is efficiently simulated by solving the incompressible Navier-Stokes equations. The solution method is based on the pseudocompressibility approach and uses an implicit-upwind differencing scheme together with the Gauss-Seidel line relaxation method. the equations are solved in steadily rotating reference frames and the centrifugal force and the Coriolis force are added to the equation of motion. Current computations use a one-equation Baldwin-Barth turbulence model which is derived from a simplified form of the standard k-epsilon model equations. The resulting computer code is applied to the flow analysis inside a generic rocket engine pump inducer, a fuel pump impeller, and SSME high pressure fuel turbopump impeller. Numerical results of inducer flow are compared with experimental measurements. In the fuel pump impeller, the effect of downstream boundary conditions is investigated. Flow analyses at 80 percent, 100 percent, and 120 percent of design conditions are presented.

SpaceX rocket fuel plan under scrutiny

NASA Astrophysics Data System (ADS)

Gwynne, Peter

2016-12-01

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

Hybrid rocket propellants from lunar material

NASA Astrophysics Data System (ADS)

Sparks, Douglas R.

This paper examines the use of lunar material for hybrid rocket propellants. Liquid oxygen is identified as the primary oxidizer and metals such as aluminum, magnesium, calcium, titanium and silicon are compared as possible fuels. Due to the reduced transportation costs, the use of lunar materials for both oxidizer and fuel will dramatically reduce the cost of a sustained space program. The advantage of hybrid rocket systems over liquid and solid rockets is discussed. It is pointed out that this type of hybrid rocket propellant could also be obtained from asteroidal and planetary soils, thereby facilitating the exploration and industrialization of the inner solar system.

Small-Scale Hybrid Rocket Test Stand & Characterization of Swirl Injectors

NASA Astrophysics Data System (ADS)

Summers, Matt H.

Derived from the necessity to increase testing capabilities of hybrid rocket motor (HRM) propulsion systems for Daedalus Astronautics at Arizona State University, a small-scale motor and test stand were designed and developed to characterize all components of the system. The motor is designed for simple integration and setup, such that both the forward-end enclosure and end cap can be easily removed for rapid integration of components during testing. Each of the components of the motor is removable allowing for a broad range of testing capabilities. While examining injectors and their potential it is thought ideal to obtain the highest regression rates and overall motor performance possible. The oxidizer and fuel are N2O and hydroxyl-terminated polybutadiene (HTPB), respectively, due to previous experience and simplicity. The injector designs, selected for the same reasons, are designed such that they vary only in the swirl angle. This system provides the platform for characterizing the effects of varying said swirl angle on HRM performance.

Solid rocket combustion simulator technology using the hybrid rocket for simulation

NASA Technical Reports Server (NTRS)

Ramohalli, Kumar

1994-01-01

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

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.

Paraffin-based hybrid rocket engines applications: A review and a market perspective

NASA Astrophysics Data System (ADS)

Mazzetti, Alessandro; Merotto, Laura; Pinarello, Giordano

2016-09-01

Hybrid propulsion technology for aerospace applications has received growing attention in recent years due to its important advantages over competitive solutions. Hybrid rocket engines have a great potential for several aeronautics and aerospace applications because of their safety, reliability, low cost and high performance. As a consequence, this propulsion technology is feasible for a number of innovative missions, including space tourism. On the other hand, hybrid rocket propulsion's main drawback, i.e. the difficulty in reaching high regression rate values using standard fuels, has so far limited the maturity level of this technology. The complex physico-chemical processes involved in hybrid rocket engines combustion are of major importance for engine performance prediction and control. Therefore, further investigation is ongoing in order to achieve a more complete understanding of such phenomena. It is well known that one of the most promising solutions for overcoming hybrid rocket engines performance limits is the use of liquefying fuels. Such fuels can lead to notably increased solid fuel regression rate due to the so-called "entrainment phenomenon". Among liquefying fuels, paraffin-based formulations have great potentials as solid fuels due to their low cost, availability (as they can be derived from industrial waste), low environmental impact and high performance. Despite the vast amount of literature available on this subject, a precise focus on market potential of paraffins for hybrid propulsion aerospace applications is lacking. In this work a review of hybrid rocket engines state of the art was performed, together with a detailed analysis of the possible applications of such a technology. A market study was carried out in order to define the near-future foreseeable development needs for hybrid technology application to the aforementioned missions. Paraffin-based fuels are taken into account as the most promising segment for market development.The present study is useful for driving future investigation and testing of paraffin-based fuels as solid fuels for hybrid propulsion technology, taking into account the needs of industrial applications of this technology.

NASA Engineer Examines the Design of a Regeneratively-Cooled Rocket Engine

NASA Image and Video Library

1958-12-21

An engineer at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a drawing showing the assembly and details of a 20,000-pound thrust regeneratively cooled rocket engine. The engine was being designed for testing in Lewis’ new Rocket Engine Test Facility, which began operating in the fall of 1957. The facility was the largest high-energy test facility in the country that was capable of handling liquid hydrogen and other liquid chemical fuels. The facility’s use of subscale engines up to 20,000 pounds of thrust permitted a cost-effective method of testing engines under various conditions. The Rocket Engine Test Facility was critical to the development of the technology that led to the use of hydrogen as a rocket fuel and the development of lightweight, regeneratively-cooled, hydrogen-fueled rocket engines. Regeneratively-cooled engines use the cryogenic liquid hydrogen as both the propellant and the coolant to prevent the engine from burning up. The fuel was fed through rows of narrow tubes that surrounded the combustion chamber and nozzle before being ignited inside the combustion chamber. The tubes are visible in the liner sitting on the desk. At the time, Pratt and Whitney was designing a 20,000-pound thrust liquid-hydrogen rocket engine, the RL-10. Two RL-10s would be used to power the Centaur second-stage rocket in the 1960s. The successful development of the Centaur rocket and the upper stages of the Saturn V were largely credited to the work carried out Lewis.

History of Sulphur Content Effects on the Thermal Stability of RP-1 under Heated Conditions

NASA Technical Reports Server (NTRS)

Irvine, Solveig A.; Schoettmer, Amanda K.; Bates, Ronald W.; Meyer, Michael L.

2004-01-01

As technologies advance in the aerospace industry, a strong desire has emerged to design more efficient, longer life, reusable liquid hydrocarbon fueled rocket engines. To achieve this goal, a more complete understanding of the thermal stability and chemical makeup of the hydrocarbon propellant is needed. Since the main fuel used in modern liquid hydrocarbon systems is RP-1, there is concern that Standard Grade RP-1 may not be a suitable propellant for future-generation rocket engines due to concern over the outdated Mil-Specification for the fuel. This current specification allows high valued limits on contaminants such as sulfur compounds, and also lacks specification of required thermal stability qualifications for the fuel. Previous studies have highlighted the detrimental effect of high levels of mercaptan sulfur content (^50 ppm) on copper rocket engine materials, but the fuel itself has not been studied. While the role of sulfur in other fuels (e.g., aviation, diesel, and automotive fuels) has been extensively studied, little has been reported on the effects of sulfur levels in rocket fuels. Lower RP-1 sulfur concentrations need to be evaluated and an acceptable sulfur limit established before RP-1 can be recommended for use as the propellant for future launch vehicles. (5 tables, 8 figures, 9 refs.)

Theoretical Studies of Ionic Liquids and Nanoclusters as Hybrid Fuels

DTIC Science & Technology

2016-08-17

Acknowledgements Distribution A: Approved for Public Release; Distribution Unlimited. PA# 16409 Aerospace Systems Directorate RQ-West (EAFB, CA)  Rocket ...Engines & Motors  Satellite Propulsion  Combustion Devices  Fuels and Propellants  System Analysis  R&D Rocket Testing RQ-East (WPAFB, OH)  Air...Distribution A: Approved for Public Release; Distribution Unlimited. PA# 16409 5 Identify and develop advanced chemical propellants for rocket

Making Ceramic Components For Advanced Aircraft Engines

NASA Technical Reports Server (NTRS)

Franklin, J. E.; Ezis, A.

1994-01-01

Lightweight, oxidation-resistant silicon nitride components containing intricate internal cooling and hydraulic passages and capable of withstanding high operating temperatures made by ceramic-platelet technology. Used to fabricate silicon nitride test articles of two types: components of methane-cooled regenerator for air turbo ramjet engine and components of bipropellant injector for rocket engine. Procedures for development of more complex and intricate components established. Technology has commercial utility in automotive, aircraft, and environmental industries for manufacture of high-temperature components for use in regeneration of fuels, treatment of emissions, high-temperature combustion devices, and application in which other high-temperature and/or lightweight components needed. Potential use in fabrication of combustors and high-temperature acoustic panels for suppression of noise in future high-speed aircraft.

Production Strategies for Production-Quality Parts for Aerospace Applications

NASA Technical Reports Server (NTRS)

Cawley, J. D.; Best, J. E.; Liu, Z.; Eckel, A. J.; Reed, B. D.; Fox, D. S.; Bhatt, R.; Levine, Stanley R. (Technical Monitor)

2000-01-01

A combination of rapid prototyping processes (3D Systems' stereolithography and Sanders Prototyping's ModelMaker) are combined with gelcasting to produce high quality silicon nitride components that were performance tested under simulated use conditions. Two types of aerospace components were produced, a low-force rocket thruster and a simulated airfoil section. The rocket was tested in a test stand using varying mixtures of H2 and O2, whereas the simulated airfoil was tested by subjecting it to a 0.3 Mach jet-fuel burner flame. Both parts performed successfully, demonstrating the usefulness of the rapid prototyping in efforts to effect materials substitution. In addition, the simulated airfoil was used to explore the possibility of applying thermal/environmental barrier coatings and providing for internal cooling of ceramic parts. It is concluded that this strategy for processing offers the ceramic engineer all the flexibility normally associated with investment casting of superalloys.

Space reactor fuel element testing in upgraded TREAT

NASA Astrophysics Data System (ADS)

Todosow, M.; Bezler, P.; Ludewig, H.; Kato, W. Y.

The testing of candidate fuel elements at prototypic operating conditions with respect to temperature, power density, hydrogen coolant flow rate, etc.; a crucial component in the development and qualification of nuclear rocket engines based on the Particle Bed Reactor (PBR); NERVA-derivative; and other concepts are discussed. Such testing may be performed at existing reactors, or at new facilities. A scoping study has been performed to assess the feasibility of testing PBR based fuel elements at the TREAT reactor. Initial results suggest that full-scale PBR elements could be tested at an average energy deposition of approximately 60-80 MW-s/L in the current TREAT reactor. If the TREAT reactor was upgraded to include fuel elements with a higher temperature limit, average energy deposition of approximately 100 MW/L may be achievable.

Space reactor fuel element testing in upgraded TREAT

NASA Astrophysics Data System (ADS)

Todosow, Michael; Bezler, Paul; Ludewig, Hans; Kato, Walter Y.

1993-01-01

The testing of candidate fuel elements at prototypic operating conditions with respect to temperature, power density, hydrogen coolant flow rate, etc., is a crucial component in the development and qualification of nuclear rocket engines based on the Particle Bed Reactor (PBR), NERVA-derivative, and other concepts. Such testing may be performed at existing reactors, or at new facilities. A scoping study has been performed to assess the feasibility of testing PBR based fuel elements at the TREAT reactor. Initial results suggests that full-scale PBR elements could be tested at an average energy deposition of ˜60-80 MW-s/L in the current TREAT reactor. If the TREAT reactor was upgraded to include fuel elements with a higher temperture limit, average energy deposition of ˜100 MW/L may be achievable.

Evaluation of Foam Coolants.

DTIC Science & Technology

HYPERGOLIC ROCKET PROPELLANTS, * FOAM , FILM COOLING, FILM COOLING, LIQUID COOLING, LIQUID ROCKET FUELS, ADDITIVES, HEAT TRANSFER, COOLANTS, LIQUID PROPELLANT ROCKET ENGINES, LIQUID COOLING, CAPTIVE TESTS, FEASIBILITY STUDIES.

Hybrid propulsion technology program: Phase 1, volume 4

NASA Technical Reports Server (NTRS)

Claflin, S. E.; Beckman, A. W.

1989-01-01

The use of a liquid oxidizer-solid fuel hybrid propellant combination in booster rocket motors appears extremely attractive due to the integration of the best features of liquid and solid propulsion systems. The hybrid rocket combines the high performance, clean exhaust, and safety of liquid propellant engines with the low cost and simplicity of solid propellant motors. Additionally, the hybrid rocket has unique advantages such as an inert fuel grain and a relative insensitivity to fuel grain and oxidizer injection anomalies. The advantages mark the hybrid rocket as a potential replacement or alternative for current and future solid propellant booster systems. The issues are addressed and recommendations are made concerning oxidizer feed systems, injectors, and ignition systems as related to hybrid rocket propulsion. Early in the program a baseline hybrid configuration was established in which liquid oxygen would be injected through ports in a solid fuel whose composition is based on hydroxyl terminated polybutadiene (HTPB). Liquid oxygen remained the recommended oxidizer and thus all of the injector concepts which were evaluated assumed only liquid would be used as the oxidizer.

Investigation of gaseous nuclear rocket technology

NASA Technical Reports Server (NTRS)

Kendall, J. S.

1972-01-01

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

FNAS/summer faculty fellowship research continuation program. Task 6: Integrated model development for liquid fueled rocket propulsion systems. Task 9: Aspects of model-based rocket engine condition monitoring and control

NASA Technical Reports Server (NTRS)

Santi, L. Michael; Helmicki, Arthur J.

1993-01-01

The objective of Phase I of this research effort was to develop an advanced mathematical-empirical model of SSME steady-state performance. Task 6 of Phase I is to develop component specific modification strategy for baseline case influence coefficient matrices. This report describes the background of SSME performance characteristics and provides a description of the control variable basis of three different gains models. The procedure used to establish influence coefficients for each of these three models is also described. Gains model analysis results are compared to Rocketdyne's power balance model (PBM).

Advanced rocket propulsion

NASA Technical Reports Server (NTRS)

Obrien, Charles J.

1993-01-01

Existing NASA research contracts are supporting development of advanced reinforced polymer and metal matrix composites for use in liquid rocket engines of the future. Advanced rocket propulsion concepts, such as modular platelet engines, dual-fuel dual-expander engines, and variable mixture ratio engines, require advanced materials and structures to reduce overall vehicle weight as well as address specific propulsion system problems related to elevated operating temperatures, new engine components, and unique operating processes. High performance propulsion systems with improved manufacturability and maintainability are needed for single stage to orbit vehicles and other high performance mission applications. One way to satisfy these needs is to develop a small engine which can be clustered in modules to provide required levels of total thrust. This approach should reduce development schedule and cost requirements by lowering hardware lead times and permitting the use of existing test facilities. Modular engines should also reduce operational costs associated with maintenance and parts inventories.

Blast from the past

NASA Astrophysics Data System (ADS)

Carlowicz, Michael

1996-02-01

Forget dynamite or hydraulic and mechanical drills. Industrial and federal researchers have started boring holes with rocket fuel. In a cooperative arrangement between Sandia National Laboratory, Global Environmental Solutions, and Universal Tech Corp., scientists and engineers extracted fuel from 200 rocket motors and used it as a mining explosive. In a demonstration completed last fall, researchers used 4950 kg of solid rocket propellant to move more than 22,500 metric tons of rock from the Lone Star Quarry in Prairie, Oklahoma. They found that the fuel improved blast energy and detonation velocity over traditional explosives, and it required fewer drill holes.

Hybrid rocket engine, theoretical model and experiment

NASA Astrophysics Data System (ADS)

Chelaru, Teodor-Viorel; Mingireanu, Florin

2011-06-01

The purpose of this paper is to build a theoretical model for the hybrid rocket engine/motor and to validate it using experimental results. The work approaches the main problems of the hybrid motor: the scalability, the stability/controllability of the operating parameters and the increasing of the solid fuel regression rate. At first, we focus on theoretical models for hybrid rocket motor and compare the results with already available experimental data from various research groups. A primary computation model is presented together with results from a numerical algorithm based on a computational model. We present theoretical predictions for several commercial hybrid rocket motors, having different scales and compare them with experimental measurements of those hybrid rocket motors. Next the paper focuses on tribrid rocket motor concept, which by supplementary liquid fuel injection can improve the thrust controllability. A complementary computation model is also presented to estimate regression rate increase of solid fuel doped with oxidizer. Finally, the stability of the hybrid rocket motor is investigated using Liapunov theory. Stability coefficients obtained are dependent on burning parameters while the stability and command matrixes are identified. The paper presents thoroughly the input data of the model, which ensures the reproducibility of the numerical results by independent researchers.

Fiber-reinforced ceramic composites for Earth-to-orbit rocket engine turbines

NASA Technical Reports Server (NTRS)

Brockmeyer, Jerry W.; Schnittgrund, Gary D.

1990-01-01

Fiber reinforced ceramic matrix composites (FRCMC) are emerging materials systems that offer potential for use in liquid rocket engines. Advantages of these materials in rocket engine turbomachinery include performance gain due to higher turbine inlet temperature, reduced launch costs, reduced maintenance with associated cost benefits, and reduced weight. This program was initiated to assess the state of FRCMC development and to propose a plan for their implementation into liquid rocket engine turbomachinery. A complete range of FRCMC materials was investigated relative to their development status and feasibility for use in the hot gas path of earth-to-orbit rocket engine turbomachinery. Of the candidate systems, carbon fiber-reinforced silicon carbide (C/SiC) offers the greatest near-term potential. Critical hot gas path components were identified, and the first stage inlet nozzle and turbine rotor of the fuel turbopump for the liquid oxygen/hydrogen Space Transportation Main Engine (STME) were selected for conceptual design and analysis. The critical issues associated with the use of FRCMC were identified. Turbine blades were designed, analyzed and fabricated. The Technology Development Plan, completed as Task 5 of this program, provides a course of action for resolution of these issues.

Nuclear Thermal Propulsion (NTP): A Proven Growth Technology for Human NEO/Mars Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The nuclear thermal rocket (NTR) represents the next "evolutionary step" in high performance rocket propulsion. Unlike conventional chemical rockets that produce their energy through combustion, the NTR derives its energy from fission of Uranium-235 atoms contained within fuel elements that comprise the engine s reactor core. Using an "expander" cycle for turbopump drive power, hydrogen propellant is raised to a high pressure and pumped through coolant channels in the fuel elements where it is superheated then expanded out a supersonic nozzle to generate high thrust. By using hydrogen for both the reactor coolant and propellant, the NTR can achieve specific impulse (Isp) values of 900 seconds (s) or more - twice that of today s best chemical rockets. From 1955 - 1972, twenty rocket reactors were designed, built and ground tested in the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs. These programs demonstrated: (1) high temperature carbide-based nuclear fuels; (2) a wide range of thrust levels; (3) sustained engine operation; (4) accumulated lifetime at full power; and (5) restart capability - all the requirements needed for a human Mars mission. Ceramic metal "cermet" fuel was pursued as well, as a backup option. The NTR also has significant "evolution and growth" capability. Configured as a "bimodal" system, it can generate its own electrical power to support spacecraft operational needs. Adding an oxygen "afterburner" nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, no large technology scale-ups are required for NTP either. In fact, the smallest engine tested during the Rover program - the 25,000 lbf (25 klbf) "Pewee" engine is sufficient when used in a clustered engine arrangement. The "Copernicus" crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth object (NEO) and Mars orbital missions prior to a Mars landing mission. The paper also discusses NASA s current activities and future plans for NTP development that include system-level Technology Demonstrations - specifically ground testing a small, scalable NTR by 2020, with a flight test shortly thereafter.

Application of Low Melting Point Thermoplastics to Hybrid Rocket Fuel

NASA Astrophysics Data System (ADS)

Wada, Yutaka; Jikei, Mitsutoshi; Kato, Ryuichi; Kato, Nobuji; Hori, Keiichi

This paper introduces the application of low melting point thermoplastics (LT) to hybrid rocket fuel. LT made by Katazen Corporation has an excellent mechanical property comparing with other thermoplastics and prospect of high surface regression rate because it has a similar physical property with low melting point of paraffin fuel which has high regression rate probably due to the entrainment mass transfer mechanism that droplets continuously depart out of the surface melt layer. Several different types of LT developed by Katazen Corporation for this use have been evaluated in the measurements of regression rate, mechanical properties These results show the LTs have the higher regression rate and better mechanical properties comparing with conventional hybrid rocket fuels. Observation was also made using a small 2D combustor, and the entrainment mass transfer mechanism is confirmed with the LT fuels.

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

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

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 approximately 10 cm long and are heated via direct electric resistive heating. This heating brings the gasses to their minimum required ignition temperature, which is lower than the auto-thermal ignition temperature, and causes the onset of both surface and gas phase ignition producing hot temperatures and a highly reacting flame. The combustion products from the catalytic tubes, which are below the melting point of platinum, are injected into the center of another combustion stage, called the primary augmenter. The reactants for this combustion stage come from the same source but the flows of non-premixed methane and oxygen gas are split off to a secondary mixing apparatus and can be mixed in a near-stoichiometric to highly lean mixture ratio. The primary augmenter is a component that has channels venting this mixed gas to impinge on each other in the center of the augmenter, perpendicular to the flow from the catalyst. The total crosssectional area of these channels is on a similar order as that of the catalyst. The augmenter has internal channels that act as a manifold to distribute equally the gas to the inward-venting channels. This stage creates a stable flame kernel as its flows, which are on the order of 0.01 g/s, are ignited by the combustion products of the catalyst. This stage is designed to produce combustion products in the flame kernel that exceed the autothermal ignition temperature of oxygen and methane.

Optimal dual-fuel propulsion for minimum inert weight or minimum fuel cost

NASA Technical Reports Server (NTRS)

Martin, J. A.

1973-01-01

An analytical investigation of single-stage vehicles with multiple propulsion phases has been conducted with the phasing optimized to minimize a general cost function. Some results are presented for linearized sizing relationships which indicate that single-stage-to-orbit, dual-fuel rocket vehicles can have lower inert weight than similar single-fuel rocket vehicles and that the advantage of dual-fuel vehicles can be increased if a dual-fuel engine is developed. The results also indicate that the optimum split can vary considerably with the choice of cost function to be minimized.

A research on polyether glycol replaced APCP rocket propellant

NASA Astrophysics Data System (ADS)

Lou, Tianyou; Bao, Chun Jia; Wang, Yiyang

2017-08-01

Ammonium perchlorate composite propellant (APCP) is a modern solid rocket propellant used in rocket vehicles. It differs from many traditional solid rocket propellants by the nature of how it is processed. APCP is cast into shape, as opposed to powder pressing it with black powder. This provides manufacturing regularity and repeatability, which are necessary requirements for use in the aerospace industry. For traditional APCP, ingredients normally used are ammonium peroxide, aluminum, Hydroxyl-terminated polybutadiene(HTPB), curing agency and other additives, the greatest disadvantage is that the fuel is too expensive. According to the price we collected in our country, a single kilogram of this fuel will cost 200 Yuan, which is about 35 dollars, for a fan who may use tons of the fuel in a single year, it definitely is a great deal of money. For this reason, we invented a new kind of APCP fuel. Changing adhesive agency from cross-linked htpb to cross linked polyether glycol gives a similar specific thrust, density and mechanical property while costs a lower price.

Feasibility study of palm-based fuels for hybrid rocket motor applications

NASA Astrophysics Data System (ADS)

Tarmizi Ahmad, M.; Abidin, Razali; Taha, A. Latif; Anudip, Amzaryi

2018-02-01

This paper describes the combined analysis done in pure palm-based wax that can be used as solid fuel in a hybrid rocket engine. The measurement of pure palm wax calorific value was performed using a bomb calorimeter. An experimental rocket engine and static test stand facility were established. After initial measurement and calibration, repeated procedures were performed. Instrumentation supplies carried out allow fuel regression rate measurements, oxidizer mass flow rates and stearic acid rocket motors measurements. Similar tests are also carried out with stearate acid (from palm oil by-products) dissolved with nitrocellulose and bee solution. Calculated data and experiments show that rates and regression thrust can be achieved even in pure-tested palm-based wax. Additionally, palm-based wax is mixed with beeswax characterized by higher nominal melting temperatures to increase moisturizing points to higher temperatures without affecting regression rate values. Calorie measurements and ballistic experiments were performed on this new fuel formulation. This new formulation promises driving applications in a wide range of temperatures.

Space reactor fuel element testing in upgraded TREAT

DOE Office of Scientific and Technical Information (OSTI.GOV)

Todosow, M.; Bezler, P.; Ludewig, H.

1993-01-14

The testing of candidate fuel elements at prototypic operating conditions with respect to temperature, power density, hydrogen coolant flow rate, etc., a crucial component in the development and qualification of nuclear rocket engines based on the Particle Bed Reactor (PBR), NERVA-derivative, and other concepts. Such testing may be performed at existing reactors, or at new facilities. A scoping study has been performed to assess the feasibility of testing PBR based fuel elements at the TREAT reactor. initial results suggest that full-scale PBR, elements could be tested at an average energy deposition of {approximately}60--80 MW-s/L in the current TREAT reactor. Ifmore » the TREAT reactor was upgraded to include fuel elements with a higher temperature limit, average energy deposition of {approximately}100 MW/L may be achievable.« less

Space reactor fuel element testing in upgraded TREAT

DOE Office of Scientific and Technical Information (OSTI.GOV)

Todosow, M.; Bezler, P.; Ludewig, H.

1993-05-01

The testing of candidate fuel elements at prototypic operating conditions with respect to temperature, power density, hydrogen coolant flow rate, etc., a crucial component in the development and qualification of nuclear rocket engines based on the Particle Bed Reactor (PBR), NERVA-derivative, and other concepts. Such testing may be performed at existing reactors, or at new facilities. A scoping study has been performed to assess the feasibility of testing PBR based fuel elements at the TREAT reactor. initial results suggest that full-scale PBR, elements could be tested at an average energy deposition of {approximately}60--80 MW-s/L in the current TREAT reactor. Ifmore » the TREAT reactor was upgraded to include fuel elements with a higher temperature limit, average energy deposition of {approximately}100 MW/L may be achievable.« less

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

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

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

NASA Astrophysics Data System (ADS)

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

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

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

NASA Technical Reports Server (NTRS)

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

1959-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2004-09-01

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

High-Energy Propellant Rocket Firing at the Rocket Lab

NASA Image and Video Library

1955-01-21

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

Experimental study of combustion in hydrogen peroxide hybrid rockets

NASA Astrophysics Data System (ADS)

Wernimont, Eric John

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

Graphics enhanced computer emulation for improved timing-race and fault tolerance control system analysis. [of Centaur liquid-fuel booster

NASA Technical Reports Server (NTRS)

Szatkowski, G. P.

1983-01-01

A computer simulation system has been developed for the Space Shuttle's advanced Centaur liquid fuel booster rocket, in order to conduct systems safety verification and flight operations training. This simulation utility is designed to analyze functional system behavior by integrating control avionics with mechanical and fluid elements, and is able to emulate any system operation, from simple relay logic to complex VLSI components, with wire-by-wire detail. A novel graphics data entry system offers a pseudo-wire wrap data base that can be easily updated. Visual subsystem operations can be selected and displayed in color on a six-monitor graphics processor. System timing and fault verification analyses are conducted by injecting component fault modes and min/max timing delays, and then observing system operation through a red line monitor.

4. Historic photo of fuel and oxidant tanks in hilltop ...

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

4. Historic photo of fuel and oxidant tanks in hilltop area of rocket engine test facility. 1956. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-1956-160D. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

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

NASA Technical Reports Server (NTRS)

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

2007-01-01

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

Fuel Regression Rate Behavior of CAMUI Hybrid Rocket

NASA Astrophysics Data System (ADS)

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

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

The Rocket Equation Improvement under ICF Implosion Experiment

NASA Astrophysics Data System (ADS)

Wang, Yanbin; Zheng, Zhijian

2013-10-01

The ICF explosion process has been studied in details. The rocket equation has been improved in explosive process by introducing the pressure parameter of fuel. Some methods could be drawn by the improved rocket equation. And the methods could be used to improve ICF target design, driving pulse design and experimental design. The First is to increase ablation pressure. The second is to decrease pressure of fuel. The third is to use larger diameter of target sphere. And the forth is to a shorten driving pulse.

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.

27 CFR 21.55 - Formula No. 28-A.

Code of Federal Regulations, 2010 CFR

2010-04-01

... hydrocarbon solvent. (b) Authorized uses. (1) As a fuel: 611.Automobile and supplementary fuels. 612.Airplane and supplementary fuels. 613.Rocket and jet fuels. 620.Proprietary heating fuels. 630.Other fuel uses...

27 CFR 21.55 - Formula No. 28-A.

Code of Federal Regulations, 2012 CFR

2012-04-01

... hydrocarbon solvent. (b) Authorized uses. (1) As a fuel: 611.Automobile and supplementary fuels. 612.Airplane and supplementary fuels. 613.Rocket and jet fuels. 620.Proprietary heating fuels. 630.Other fuel uses...

27 CFR 21.55 - Formula No. 28-A.

Code of Federal Regulations, 2011 CFR

2011-04-01

... hydrocarbon solvent. (b) Authorized uses. (1) As a fuel: 611.Automobile and supplementary fuels. 612.Airplane and supplementary fuels. 613.Rocket and jet fuels. 620.Proprietary heating fuels. 630.Other fuel uses...

27 CFR 21.55 - Formula No. 28-A.

Code of Federal Regulations, 2014 CFR

2014-04-01

... hydrocarbon solvent. (b) Authorized uses. (1) As a fuel: 611.Automobile and supplementary fuels. 612.Airplane and supplementary fuels. 613.Rocket and jet fuels. 620.Proprietary heating fuels. 630.Other fuel uses...

27 CFR 21.55 - Formula No. 28-A.

Code of Federal Regulations, 2013 CFR

2013-04-01

... hydrocarbon solvent. (b) Authorized uses. (1) As a fuel: 611.Automobile and supplementary fuels. 612.Airplane and supplementary fuels. 613.Rocket and jet fuels. 620.Proprietary heating fuels. 630.Other fuel uses...

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

NASA Astrophysics Data System (ADS)

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

2015-09-01

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

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.

Chromosomal instability in rodents caused by pollution from Baikonur cosmodrome.

PubMed

Kolumbayeva, Saule; Begimbetova, Dinara; Shalakhmetova, Tamara; Saliev, Timur; Lovinskaya, Anna; Zhunusbekova, Benazir

2014-09-01

An assessment of the health status of ecosystems exposed to man-made pollution is a vital issue for many countries. Particularly it concerns the consequences of contamination caused by the activity of the space industry. Each rocket launch is accompanied by the introduction of parts of the rocket propellant into the environment. This study aims to scrutinize the effect of the components of rocket fuel on the induction of lipid peroxidation and chromosomal aberrations on rodents inhabiting the area exposed to pollution from Baikonur cosmodrome. The results showed the increase of the level of lipid hydroperoxide and malondialdehyde in the livers of Citellus pygmaeus Pallas and Mus musculus L., which indicates an augmentation of free radical activity and DNA damage. The cytogenetic analysis of bone marrow cells revealed that the frequency of chromosomal aberrations was a few times higher in the rodents from contaminated territory. The signs of oxidative stress and high level of chromosomal aberrations indicate the environmental impact of the cosmodrome, and its possible toxic and mutagenic effects on ecosystems.

It is rocket science – why dietary nitrate is hard to ‘beet’! Part II: further mechanisms and therapeutic potential of the nitrate‐nitrite‐NO pathway

PubMed Central

Mills, Charlotte Elizabeth; Khatri, Jibran; Maskell, Perry; Odongerel, Chimed

2016-01-01

Dietary nitrate (found in green leafy vegetables such as rocket and in beetroot) is now recognized to be an important source of nitric oxide, via the nitrate‐nitrite‐NO pathway. Dietary nitrate confers several cardiovascular beneficial effects on blood pressure, platelets, endothelial function, mitochondrial efficiency and exercise. Having described key twists and turns in the elucidation of the pathway and the underlying mechanisms in Part I, we explore the more recent developments which have served to confirm mechanisms, extend our understanding, and discover new properties and potential therapeutic uses of the pathway in Part II. Even the established dependency on low oxygen states for bioactivation of nitrite has recently been challenged. Dietary nitrate appears to be an important component of ‘healthy diets’, such as the DASH diet to lower blood pressure and the Mediterranean diet, with its potential to lower cardiovascular risk, possibly through beneficial interactions with a range of other constituents. The World Cancer Research Foundation report strong evidence for vegetables including spinach and lettuce (high nitrate‐containing) decreasing cancer risk (mouth, pharynx, larynx, oesophagus and stomach), summarized in a ‘Nitrate‐Cancer Risk Veg‐Table’. The European Space Agency recommends that beetroot, lettuce, spinach and rocket (high‐nitrate vegetables) are grown to provide food for long‐term space missions. Nitrate, an ancient component of rocket fuel, could support sustainable crops for healthy humans. PMID:26914827

It is rocket science - why dietary nitrate is hard to 'beet'! Part II: further mechanisms and therapeutic potential of the nitrate-nitrite-NO pathway.

PubMed

Mills, Charlotte Elizabeth; Khatri, Jibran; Maskell, Perry; Odongerel, Chimed; Webb, Andrew James

2017-01-01

Dietary nitrate (found in green leafy vegetables such as rocket and in beetroot) is now recognized to be an important source of nitric oxide, via the nitrate-nitrite-NO pathway. Dietary nitrate confers several cardiovascular beneficial effects on blood pressure, platelets, endothelial function, mitochondrial efficiency and exercise. Having described key twists and turns in the elucidation of the pathway and the underlying mechanisms in Part I, we explore the more recent developments which have served to confirm mechanisms, extend our understanding, and discover new properties and potential therapeutic uses of the pathway in Part II. Even the established dependency on low oxygen states for bioactivation of nitrite has recently been challenged. Dietary nitrate appears to be an important component of 'healthy diets', such as the DASH diet to lower blood pressure and the Mediterranean diet, with its potential to lower cardiovascular risk, possibly through beneficial interactions with a range of other constituents. The World Cancer Research Foundation report strong evidence for vegetables including spinach and lettuce (high nitrate-containing) decreasing cancer risk (mouth, pharynx, larynx, oesophagus and stomach), summarized in a 'Nitrate-Cancer Risk Veg-Table'. The European Space Agency recommends that beetroot, lettuce, spinach and rocket (high-nitrate vegetables) are grown to provide food for long-term space missions. Nitrate, an ancient component of rocket fuel, could support sustainable crops for healthy humans. © 2016 The British Pharmacological Society.

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

NASA Astrophysics Data System (ADS)

Boronowsky, Kenny

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

Controlling Gas-Flow Mass Ratios

NASA Technical Reports Server (NTRS)

Morris, Brian G.

1990-01-01

Proposed system automatically controls proportions of gases flowing in supply lines. Conceived for control of oxidizer-to-fuel ratio in new gaseous-propellant rocket engines. Gas-flow control system measures temperatures and pressures at various points. From data, calculates control voltages for electronic pressure regulators for oxygen and hydrogen. System includes commercially available components. Applicable to control of mass ratios in such gaseous industrial processes as chemical-vapor depostion of semiconductor materials and in automotive engines operating on compressed natural gas.

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

NASA Astrophysics Data System (ADS)

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

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

Combustion performance and scale effect from N2O/HTPB hybrid rocket motor simulations

NASA Astrophysics Data System (ADS)

Shan, Fanli; Hou, Lingyun; Piao, Ying

2013-04-01

HRM code for the simulation of N2O/HTPB hybrid rocket motor operation and scale effect analysis has been developed. This code can be used to calculate motor thrust and distributions of physical properties inside the combustion chamber and nozzle during the operational phase by solving the unsteady Navier-Stokes equations using a corrected compressible difference scheme and a two-step, five species combustion model. A dynamic fuel surface regression technique and a two-step calculation method together with the gas-solid coupling are applied in the calculation of fuel regression and the determination of combustion chamber wall profile as fuel regresses. Both the calculated motor thrust from start-up to shut-down mode and the combustion chamber wall profile after motor operation are in good agreements with experimental data. The fuel regression rate equation and the relation between fuel regression rate and axial distance have been derived. Analysis of results suggests improvements in combustion performance to the current hybrid rocket motor design and explains scale effects in the variation of fuel regression rate with combustion chamber diameter.

KSC-05pd2558

NASA Image and Video Library

2005-12-05

KENNEDY SPACE CENTER, FLA. - The Lockheed Martin Atlas V rocket (center) undergoes a tanking test on Launch Complex 41 at Cape Canaveral Air Force Station in Florida. The rocket was fully fueled with liquid hydrogen, liquid oxygen and RP 1 kerosene fuel. Seen surrounding the rocket are lightning towers that support the catenary wire that provides lightning protection. The Atlas V is the launch vehicle for NASA’s New Horizons spacecraft, scheduled to launch during a 35-day window that opens Jan. 11, and fly through the Pluto system as early as summer 2015.

KSC-05pd2559

NASA Image and Video Library

2005-12-05

KENNEDY SPACE CENTER, FLA. - The Lockheed Martin Atlas V rocket (center) undergoes a tanking test on Launch Complex 41 at Cape Canaveral Air Force Station in Florida. The rocket was fully fueled with liquid hydrogen, liquid oxygen and RP 1 kerosene fuel. Seen surrounding the rocket are lightning towers that support the catenary wire that provides lightning protection. The Atlas V is the launch vehicle for NASA’s New Horizons spacecraft, scheduled to launch during a 35-day window that opens Jan. 11, and fly through the Pluto system as early as summer 2015.

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

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

Nuclear Thermal Rocket (Ntr) Propulsion: A Proven Game-Changing Technology for Future Human Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The NTR represents the next evolutionary step in high performance rocket propulsion. It generates high thrust and has a specific impulse (Isp) of approx.900 seconds (s) or more V twice that of today s best chemical rockets. The technology is also proven. During the previous Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) nuclear rocket programs, 20 rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability V all the requirements needed for a human mission to Mars. Ceramic metal cermet fuel was also pursued, as a backup option. The NTR also has significant growth and evolution potential. Configured as a bimodal system, it can generate electrical power for the spacecraft. Adding an oxygen afterburner nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, simple assembly and mission operations. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program V the 25,000 lbf (25 klbf) Pewee engine is sufficient for human Mars missions when used in a clustered engine arrangement. The Copernicus crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth asteroid (NEA) and Mars orbital missions prior to a Mars landing mission. Initially, the basic Copernicus vehicle can enable reusable 1-year round trip human missions to candidate NEAs like 1991 JW and Apophis in the late 2020 s to check out vehicle systems. Afterwards, the Copernicus spacecraft and its 2 key components, now configured as an Earth Return Vehicle / propellant tanker, would be used for a short round trip (approx.18 - 20 months)/short orbital stay (60 days) Mars / Phobos survey mission in 2033 using a split mission approach. The paper also discusses NASA s current Foundational Technology Development activities and its pre-decisional plans for future system-level Technology Demonstrations that include ground testing a small (approx.7.5 klbf) scalable NTR before the decade is out with a flight test shortly thereafter.

Materials for Liquid Propulsion Systems. Chapter 12

NASA Technical Reports Server (NTRS)

Halchak, John A.; Cannon, James L.; Brown, Corey

2016-01-01

Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.

How High? How Fast? How Long? Modeling Water Rocket Flight with Calculus

ERIC Educational Resources Information Center

Ashline, George; Ellis-Monaghan, Joanna

2006-01-01

We describe an easy and fun project using water rockets to demonstrate applications of single variable calculus concepts. We provide procedures and a supplies list for launching and videotaping a water rocket flight to provide the experimental data. Because of factors such as fuel expulsion and wind effects, the water rocket does not follow the…

Reactive Shear Layer Mixing and Growth Rate Effects on Afterburning Properties for Axisymetric Rocket Engine Plumes

DTIC Science & Technology

2006-09-01

water, carbon monoxide and carbon dioxide . The ratio of specific heats is reduced as the number of atoms in the molecule increases and as the...The flow of the jet is faster than the surrounding air, and since gas turbine engines run fuel lean, the exhaust products have generally fully reacted...previous types by several characteristics. The core of the rocket exhaust flowfield is fuel rich, and unlike gas turbine engines, which burn fuel

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).

The optimisation of low-acceleration interstellar relativistic rocket trajectories using genetic algorithms

NASA Astrophysics Data System (ADS)

Fung, Kenneth K. H.; Lewis, Geraint F.; Wu, Xiaofeng

2017-04-01

A vast wealth of literature exists on the topic of rocket trajectory optimisation, particularly in the area of interplanetary trajectories due to its relevance today. Studies on optimising interstellar and intergalactic trajectories are usually performed in flat spacetime using an analytical approach, with very little focus on optimising interstellar trajectories in a general relativistic framework. This paper examines the use of low-acceleration rockets to reach galactic destinations in the least possible time, with a genetic algorithm being employed for the optimisation process. The fuel required for each journey was calculated for various types of propulsion systems to determine the viability of low-acceleration rockets to colonise the Milky Way. The results showed that to limit the amount of fuel carried on board, an antimatter propulsion system would likely be the minimum technological requirement to reach star systems tens of thousands of light years away. However, using a low-acceleration rocket would require several hundreds of thousands of years to reach these star systems, with minimal time dilation effects since maximum velocities only reached about 0.2 c . Such transit times are clearly impractical, and thus, any kind of colonisation using low acceleration rockets would be difficult. High accelerations, on the order of 1 g, are likely required to complete interstellar journeys within a reasonable time frame, though they may require prohibitively large amounts of fuel. So for now, it appears that humanity's ultimate goal of a galactic empire may only be possible at significantly higher accelerations, though the propulsion technology requirement for a journey that uses realistic amounts of fuel remains to be determined.

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

NASA Technical Reports Server (NTRS)

Park, C.

1976-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2003-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

Performance estimates for space shuttle vehicles using a hydrogen or a methane fueled turboramjet powered first stage

NASA Technical Reports Server (NTRS)

Knip, G., Jr.; Eisenberg, J. D.

1972-01-01

Two- and three-stage (second stage expendable) shuttle vehicles, both having a hydrogen-fueled, turboramjet-powered first stage, are compared with a two-stage, VTOHL, all-rocket shuttle in terms of payload fraction, inert weight, development cost, operating cost, and total cost. All of the vehicles place 22,680 kilograms of payload into a 500-kilometer orbit. The upper stage(s) uses hydrogen-oxygen rockets. The effect on payload fraction and vehicle inert weight of methane and methane-FLOX as a fuel-propellant combination for the three-stage vehicle is indicated. Compared with a rocket first stage for a two-stage shuttle, an airbreathing first stage results in a higher payload fraction and a lower operating cost, but a higher total cost. The effect on cost of program size and first-stage flyback is indicated. The addition of an expendable rocket second stage (three-stage vehicle) improves the payload fraction but is unattractive economically.

Pad Safety Personnel Launch Support For STS-200

NASA Technical Reports Server (NTRS)

Guarino, Jennifer

2007-01-01

The launch of a space shuttle is a complex and lengthy procedure. There are many places and components to look at and prepare. The components are the orbiter, solid rocket boosters, external tank, and ground equipment. Some of the places are the launch pad, fuel locations, and surrounding structures. Preparations for a launch include equipment checks, system checks, sniff checks for hazardous commodities, and countless walkdowns. Throughout these preparations, pad safety personnel must always be on call. This requires three shifts of multiple people to be ready when needed. Also, the pad safety personnel must be available for the non-launch tasks that are always present for both launch pads

Researcher Poses with a Nuclear Rocket Model

NASA Image and Video Library

1961-11-21

A researcher at the NASA Lewis Research Center with slide ruler poses with models of the earth and a nuclear-propelled rocket. The Nuclear Engine for Rocket Vehicle Applications (NERVA) was a joint NASA and Atomic Energy Commission (AEC) endeavor to develop a nuclear-powered rocket for both long-range missions to Mars and as a possible upper-stage for the Apollo Program. The early portion of the program consisted of basic reactor and fuel system research. This was followed by a series of Kiwi reactors built to test nuclear rocket principles in a non-flying nuclear engine. The next phase, NERVA, would create an entire flyable engine. The AEC was responsible for designing the nuclear reactor and overall engine. NASA Lewis was responsible for developing the liquid-hydrogen fuel system. The nuclear rocket model in this photograph includes a reactor at the far right with a hydrogen propellant tank and large radiator below. The payload or crew would be at the far left, distanced from the reactor.

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.

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.

Iridium-Coated Rhenium Radiation-Cooled Rockets

NASA Technical Reports Server (NTRS)

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

1997-01-01

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

Piping Connector

NASA Technical Reports Server (NTRS)

1994-01-01

In Stennis Space Center's Component Test Facility, piping lines carry rocket propellants and high pressure cryogenic fuels. When the lines are chilled to a pretest temperature of 400 degrees below zero, ordinary piping connectors can leak. Under contract to Stennis, Reflange, Inc. developed the T-Con connector, which included a secondary seal that tolerates severe temperature change. Because of the limited need for the large and expensive T-Con product, Reflange also developed the less costly E-Con, a smaller more compact design with the same technical advantages as the T-Con.

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.

Design of a Hybrid Propulsion System for Orbit Raising Applications

NASA Astrophysics Data System (ADS)

Boman, N.; Ford, M.

2004-10-01

A trade off between conventional liquid apogee engines used for orbit raising applications and hybrid rocket engines (HRE) has been performed using a case study approach. Current requirements for lower cost and enhanced safety places hybrid propulsion systems in the spotlight. For evaluating and design of a hybrid rocket engine a parametric engineering code is developed, based on the combustion chamber characteristics of selected propellants. A single port cylindrical section of fuel grain is considered. Polyethylene (PE) and hydroxyl-terminated polybutadiene (HTPB) represents the fuels investigated. The engine design is optimized to minimize the propulsion system volume and mass, while keeping the system as simple as possible. It is found that the fuel grain L/D ratio boundary condition has a major impact on the overall hybrid rocket engine design.

Acoustic effects of sprays

NASA Technical Reports Server (NTRS)

Pindera, Maciej Z.; Przekwas, Andrzej J.

1994-01-01

Since the early 1960's, it has been known that realistic combustion models for liquid fuel rocket engines should contain at least a rudimentary treatment of atomization and spray physics. This is of particular importance in transient operations. It has long been recognized that spray characteristics and droplet vaporization physics play a fundamental role in determining the stability behavior of liquid fuel rocket motors. This paper gives an overview of work in progress on design of a numerical algorithm for practical studies of combustion instabilities in liquid rocket motors. For flexibility, the algorithm is composed of semi-independent solution modules, accounting for different physical processes. Current findings are report and future work is indicated. The main emphasis of this research is the development of an efficient treatment to interactions between acoustic fields and liquid fuel/oxidizer sprays.

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.

Space Shuttle Projects

NASA Image and Video Library

1975-01-01

As early as September 1972, the Marshall Space Flight Center arnounced plans for a series of 20 water-entry simulation tests with a solid-fueled rocket casing assembly. The tests would provide valuable data for assessment of solid rocket booster parachute water recovery and aid in preliminary solid rocket motor design.

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

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

Numerical analysis of combustion characteristics of hybrid rocket motor with multi-section swirl injection

NASA Astrophysics Data System (ADS)

Li, Chengen; Cai, Guobiao; Tian, Hui

2016-06-01

This paper is aimed to analyse the combustion characteristics of hybrid rocket motor with multi-section swirl injection by simulating the combustion flow field. Numerical combustion flow field and combustion performance parameters are obtained through three-dimensional numerical simulations based on a steady numerical model proposed in this paper. The hybrid rocket motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Multiple injection sections are set along the axis of the solid fuel grain, and the oxidizer enters the combustion chamber by means of tangential injection via the injector ports in the injection sections. Simulation results indicate that the combustion flow field structure of the hybrid rocket motor could be improved by multi-section swirl injection method. The transformation of the combustion flow field can greatly increase the fuel regression rate and the combustion efficiency. The average fuel regression rate of the motor with multi-section swirl injection is improved by 8.37 times compared with that of the motor with conventional head-end irrotational injection. The combustion efficiency is increased to 95.73%. Besides, the simulation results also indicate that (1) the additional injection sections can increase the fuel regression rate and the combustion efficiency; (2) the upstream offset of the injection sections reduces the combustion efficiency; and (3) the fuel regression rate and the combustion efficiency decrease with the reduction of the number of injector ports in each injection section.

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

NASA Technical Reports Server (NTRS)

Putre, H. A.

1971-01-01

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

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

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.

2017-01-01

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

Numerical investigations of hybrid rocket engines

NASA Astrophysics Data System (ADS)

Betelin, V. B.; Kushnirenko, A. G.; Smirnov, N. N.; Nikitin, V. F.; Tyurenkova, V. V.; Stamov, L. I.

2018-03-01

Paper presents the results of numerical studies of hybrid rocket engines operating cycle including unsteady-state transition stage. A mathematical model is developed accounting for the peculiarities of diffusion combustion of fuel in the flow of oxidant, which is composed of oxygen-nitrogen mixture. Three dimensional unsteady-state simulations of chemically reacting gas mixture above thermochemically destructing surface are performed. The results show that the diffusion combustion brings to strongly non-uniform fuel mass regression rate in the flow direction. Diffusive deceleration of chemical reaction brings to the decrease of fuel regression rate in the longitudinal direction.

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

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2008-01-01

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

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.

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.

The techniques of quality operations computational and experimental researches of the launch vehicles in the drawing-board stage

NASA Astrophysics Data System (ADS)

Rozhaeva, K.

2018-01-01

The aim of the researchis the quality operations of the design process at the stage of research works on the development of active on-Board system of the launch vehicles spent stages descent with liquid propellant rocket engines by simulating the gasification process of undeveloped residues of fuel in the tanks. The design techniques of the gasification process of liquid rocket propellant components residues in the tank to the expense of finding and fixing errors in the algorithm calculation to increase the accuracy of calculation results is proposed. Experimental modelling of the model liquid evaporation in a limited reservoir of the experimental stand, allowing due to the false measurements rejection based on given criteria and detected faults to enhance the results reliability of the experimental studies; to reduce the experiments cost.

Photoignition Torch Applied to Cryogenic H2/O2 Coaxial Jet

DTIC Science & Technology

2016-12-06

suitable for certain thrusters and liquid rocket engines. This ignition system is scalable for applications in different combustion chambers such as gas ...turbines, gas generators, liquid rocket engines, and multi grain solid rocket motors. photoignition, fuel spray ignition, high pressure ignition...thrusters and liquid rocket engines. This ignition system is scalable for applications in different combustion chambers such as gas turbines, gas

Hidden Figures To Modern Figures- Students See SLS Rocket at Michoud

NASA Image and Video Library

2017-01-09

New Orleans teacher Katherine Michelle Sanders of St. Peter Claver School, takes her 7th grade science on a tour of nearby NASA’s Michoud Assembly Facility to see where the Space Launch System - the world’s most powerful rocket - is being built. Sanders is the granddaughter of famed NASA scientist Katherine Johnson who was featured in the book and movie, Hidden Figures. Sanders discusses how her grandmother’s work inspired her and why she wanted her students to learn about NASA’s plan to explore deep space. Sanders followed her grandmother's footsteps, as Katherine Johnson was also a teacher before going to work for the National Advisory Committee for Aeronautics (NACA) and later at NASA's Langley Research Center as a human computer. Due to Johnson’s historical role as one of the first African-American women to work as a NASA scientist, she was awarded the Presidential Medal of Freedom by President Barack Obama in 2015. In this video, Michoud Director Bobby Watkins and NASA engineer Renee Horton speak with the students about key components building the SLS rocket, including the Vertical Assembly Center - a large robotic tool for manufacturing hardware - and a liquid hydrogen tank that will hold fuel for the first mission, and a liquid hydrogen tank that will hold fuel for the first SLS mission. For more information on SLS, visit www.nasa.gov/sls. For more information on Michoud Assembly Facility, visit www.nasa.gov/centers/marshall/michoud/index.html

Benefit from NASA

NASA Image and Video Library

1999-01-01

The same rocket fuel that helps power the Space Shuttle as it thunders into orbit will now be taking on a new role, with the potential to benefit millions of people worldwide. Leftover rocket fuel from NASA is being used to make a flare that destroys land mines where they were buried, without using explosives. The flare is safe to handle and easy to use. People working to deactivate the mines simply place the flare next to the uncovered land mine and ignite it from a safe distance using a battery-triggered electric match. The flare burns a hole in the land mine's case and ignites its explosive contents. The explosive burns away, disabling the mine and rendering it harmless. Using leftover rocket fuel to help destroy land mines incurs no additional costs to taxpayers. To ensure enough propellant is available for each Shuttle mission, NASA allows for a small percentage of extra propellant in each batch. Once mixed, surplus fuel solidifies and carnot be saved for use in another launch. In its solid form, it is an ideal ingredient for the new flare. The flare was developed by Thiokol Propulsion in Brigham City, Utah, the NASA contractor that designs and builds rocket motors for the Solid Rocket Booster Space Shuttle. An estimated 80 million or more active land mines are scattered around the world in at least 70 countries, and kill or maim 26,000 people a year. Worldwide, there is one casualty every 22 minutes

Land Mines Removal

NASA Technical Reports Server (NTRS)

1999-01-01

The same rocket fuel that helps power the Space Shuttle as it thunders into orbit will now be taking on a new role, with the potential to benefit millions of people worldwide. Leftover rocket fuel from NASA is being used to make a flare that destroys land mines where they were buried, without using explosives. The flare is safe to handle and easy to use. People working to deactivate the mines simply place the flare next to the uncovered land mine and ignite it from a safe distance using a battery-triggered electric match. The flare burns a hole in the land mine's case and ignites its explosive contents. The explosive burns away, disabling the mine and rendering it harmless. Using leftover rocket fuel to help destroy land mines incurs no additional costs to taxpayers. To ensure enough propellant is available for each Shuttle mission, NASA allows for a small percentage of extra propellant in each batch. Once mixed, surplus fuel solidifies and carnot be saved for use in another launch. In its solid form, it is an ideal ingredient for new the flare. The flare was developed by Thiokol Propulsion in Brigham City, Utah, the NASA contractor that designs and builds rocket motors for the Solid Rocket Booster Space Shuttle. An estimated 80 million or more active land mines are scattered around the world in at least 70 countries, and kill or maim 26,000 people a year. Worldwide, there is one casualty every 22 minutes.

Land Mines Removal

NASA Technical Reports Server (NTRS)

1999-01-01

The same rocket fuel that helps power the Space Shuttle as it thunders into orbit will now be taking on a new role, with the potential to benefit millions of people worldwide. Leftover rocket fuel from NASA is being used to make a flare that destroys land mines where they were buried, without using explosives. The flare is safe to handle and easy to use. People working to deactivate the mines simply place the flare next to the uncovered land mine and ignite it from a safe distance using a battery-triggered electric match. The flare burns a hole in the land mine's case and ignites its explosive contents. The explosive burns away, disabling the mine and rendering it harmless. Using leftover rocket fuel to help destroy land mines incurs no additional costs to taxpayers. To ensure enough propellant is available for each Shuttle mission, NASA allows for a small percentage of extra propellant in each batch. Once mixed, surplus fuel solidifies and carnot be saved for use in another launch. In its solid form, it is an ideal ingredient for the new flare. The flare was developed by Thiokol Propulsion in Brigham City, Utah, the NASA contractor that designs and builds rocket motors for the Solid Rocket Booster Space Shuttle. An estimated 80 million or more active land mines are scattered around the world in at least 70 countries, and kill or maim 26,000 people a year. Worldwide, there is one casualty every 22 minutes

Transpiration cooled throat for hydrocarbon rocket engines

NASA Technical Reports Server (NTRS)

May, Lee R.; Burkhardt, Wendel M.

1991-01-01

The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.

Hybrid Rocket Performance Prediction with Coupling Method of CFD and Thermal Conduction Calculation

NASA Astrophysics Data System (ADS)

Funami, Yuki; Shimada, Toru

The final purpose of this study is to develop a design tool for hybrid rocket engines. This tool is a computer code which will be used in order to investigate rocket performance characteristics and unsteady phenomena lasting through the burning time, such as fuel regression or combustion oscillation. When phenomena inside a combustion chamber, namely boundary layer combustion, are described, it is difficult to use rigorous models for this target. It is because calculation cost may be too expensive. Therefore simple models are required for this calculation. In this study, quasi-one-dimensional compressible Euler equations for flowfields inside a chamber and the equation for thermal conduction inside a solid fuel are numerically solved. The energy balance equation at the solid fuel surface is solved to estimate fuel regression rate. Heat feedback model is Karabeyoglu's model dependent on total mass flux. Combustion model is global single step reaction model for 4 chemical species or chemical equilibrium model for 9 chemical species. As a first step, steady-state solutions are reported.

Deposit formation in hydrocarbon rocket fuels

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

Thrust Augmented Nozzle for a Hybrid Rocket with a Helical Fuel Port

NASA Astrophysics Data System (ADS)

Marshall, Joel H.

A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor's performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

Grease-Resistant O Rings for Joints in Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Harvey, Albert R.; Feldman, Harold

2003-01-01

There is a continuing effort to develop improved O rings for sealing joints in solid-fuel rocket motors. Following an approach based on the lessons learned in the explosion of the space shuttle Challenger, investigators have been seeking O-ring materials that exhibit adequate resilience for effective sealing over a broad temperature range: What are desired are O rings that expand far and fast enough to maintain seals, even when metal sealing surfaces at a joint move slightly away from each other shortly after ignition and the motor was exposed to cold weather before ignition. Other qualities desired of the improved O rings include adequate resistance to ablation by hot rocket gases and resistance to swelling when exposed to hydrocarbon-based greases used to protect some motor components against corrosion. Five rubber formulations two based on a fluorosilicone polymer and three based on copolymers of epichlorohydrin with ethylene oxide were tested as candidate O-ring materials. Of these, one of the epichlorohydrin/ethylene oxide formulations was found to offer the closest to the desired combination of properties and was selected for further evaluation.

Ablation study of tungsten-based nuclear thermal rocket fuel

NASA Astrophysics Data System (ADS)

Smith, Tabitha Elizabeth Rose

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

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.

Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

NASA Technical Reports Server (NTRS)

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

2011-01-01

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

Feasibility of rotating fluidized bed reactor for rocket propulsion

NASA Technical Reports Server (NTRS)

Ludewig, H.; Manning, A. J.; Raseman, C. J.

1974-01-01

The rotating fluidized bed reactor concept is outlined, and its application to rocket propulsion is discussed. Experimental results obtained indicate that minimum fluidization correlations commonly in use for 1-g beds can also be applied to multiple-g beds. It was found that for a low thrust system (20,000 lbf) the fuel particle size and/or particle stress play a limiting role on performance. The superiority of U-233 as a fuel for this type of rocket engine is clearly demonstrated in the analysis. The maximum thrust/weight ratio for a 90,000N thrust engine was found to be approximately 65N/kg.

Investigating the effects of critical phenomena in premixed methane-oxygen flames at cryogenic conditions

NASA Astrophysics Data System (ADS)

Gopal, Abishek; Yellapantula, Shashank; Larsson, Johan

2017-11-01

Methane is increasingly becoming viable as a rocket fuel in the latest generation of launch vehicles. In liquid rocket engines, fuel and oxidizer are injected under cryogenic conditions into the combustion chamber. At high pressures, typical of rocket combustion chambers, the propellants exist in supercritical states where the ideal gas thermodynamics are no longer valid. We investigate the effects of real-gas thermodynamics on transcritical laminar premixed methane-oxygen flames. The effect of the real-gas cubic equations of state and high-pressure transport properties on flame dynamics is presented. We also study real-gas effects on the extinction limits of the methane-oxygen flame.

The hard start phenomena in hypergolic engines. Volume 1: Bibliography

NASA Technical Reports Server (NTRS)

Miron, Y.; Perlee, H. E.

1974-01-01

A bibliography of reports pertaining to the hard start phenomenon in attitude control rocket engines on Apollo spacecraft is presented. Some of the subjects discussed are; (1) combustion of hydrazine, (2) one dimensional theory of liquid fuel rocket combustion, (3) preignition phenomena in small pulsed rocket engines, (4) experimental and theoretical investigation of the fluid dynamics of rocket combustion, and (5) nonequilibrium combustion and nozzle flow in propellant performance.

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)

Thermophysics Characterization of Kerosene Combustion

NASA Technical Reports Server (NTRS)

Wang, Ten-See

2000-01-01

A one-formula surrogate fuel formulation and its quasi-global combustion kinetics model are developed to support the design of injectors and thrust chambers of kerosene-fueled rocket engines. This surrogate fuel model depicts a fuel blend that properly represents the general physical and chemical properties of kerosene. The accompanying gaseous-phase thermodynamics of the surrogate fuel is anchored with the heat of formation of kerosene and verified by comparing a series of one-dimensional rocket thrust chamber calculations. The quasi-global combustion kinetics model consists of several global steps for parent fuel decomposition, soot formation, and soot oxidation, and a detailed wet-CO mechanism. The final thermophysics formulations are incorporated with a computational fluid dynamics model for prediction of the combustor efficiency of an uni-element, tri-propellant combustor and the radiation of a kerosene-fueled thruster plume. The model predictions agreed reasonably well with those of the tests.

RP-1 and JP-8 Thermal Stability Experiments

NASA Technical Reports Server (NTRS)

Brown, Sarah P.; Emens, Jessica M.; Frederick, Robert A., Jr.

2005-01-01

This work experimentally investigates the effect of fuel composition changes on jet and rocket fuel thermal stability. A High Reynolds Number Thermal Stability test device evaluated JP-8 and RP-1 fuels. The experiment consisted of an electrically heated, stainless steel capillary tube with a controlled fuel outlet temperature. An optical pyrometer monitored the increasing external temperature profiles of the capillary tube as deposits build inside during each test. Multiple runs of each fuel composition provided results on measurement repeatability. Testing a t two different facilities provided data on measurement reproducibility. The technique is able to distinguish between thermally stable and unstable compositions of JP-8 and intermediate blends made by combining each composition. The technique is also able to distinguish among standard RP-1 rocket fuels and those having reduced sulfur levels. Carbon burn off analysis of residue in the capillary tubes on the RP-1 fuels correlates with the external temperature results.

Thermophysics Characterization of Kerosene Combustion

NASA Technical Reports Server (NTRS)

Wang, Ten-See

2001-01-01

A one-formula surrogate fuel formulation and its quasi-global combustion kinetics model are developed to support the design of injectors and thrust chambers of kerosene-fueled rocket engines. This surrogate fuel model depicts a fuel blend that properly represents the general physical and chemical properties of kerosene. The accompanying gaseous-phase thermodynamics of the surrogate fuel is anchored with the heat of formation of kerosene and verified by comparing a series of one-dimensional rocket thrust chamber calculations. The quasi-global combustion kinetics model consists of several global steps for parent fuel decomposition, soot formation, and soot oxidation and a detailed wet-CO mechanism to complete the combustion process. The final thermophysics formulations are incorporated with a computational fluid dynamics model for prediction of the combustion efficiency of an unielement, tripropellant combustor and the radiation of a kerosene-fueled thruster plume. The model predictions agreed reasonably well with those of the tests.

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.

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

NASA Technical Reports Server (NTRS)

Heidmann, M F

1957-01-01

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

Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

NASA Technical Reports Server (NTRS)

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

2011-01-01

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

An improved heat transfer configuration for a solid-core nuclear thermal rocket engine

NASA Technical Reports Server (NTRS)

Clark, John S.; Walton, James T.; Mcguire, Melissa L.

1992-01-01

Interrupted flow, impingement cooling, and axial power distribution are employed to enhance the heat-transfer configuration of a solid-core nuclear thermal rocket engine. Impingement cooling is introduced to increase the local heat-transfer coefficients between the reactor material and the coolants. Increased fuel loading is used at the inlet end of the reactor to enhance heat-transfer capability where the temperature differences are the greatest. A thermal-hydraulics computer program for an unfueled NERVA reactor core is employed to analyze the proposed configuration with attention given to uniform fuel loading, number of channels through the impingement wafers, fuel-element length, mass-flow rate, and wafer gap. The impingement wafer concept (IWC) is shown to have heat-transfer characteristics that are better than those of the NERVA-derived reactor at 2500 K. The IWC concept is argued to be an effective heat-transfer configuration for solid-core nuclear thermal rocket engines.

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

NASA Astrophysics Data System (ADS)

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

2014-11-01

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

U.S. Strategic Nuclear Forces: Background, Developments, and Issues

DTIC Science & Technology

2016-09-27

meet the terms of the New START Treaty. The Air Force is also modernizing the Minuteman missiles, replacing and upgrading their rocket motors...began in 1998 and has been replacing the propellant, the solid rocket fuel, in the Minuteman motors to extend the life of the rocket motors. A...complete the program. It has not requested additional funding in subsequent years. Propulsion System Rocket Engine Program (PSRE) According to the Air

Research Technology

NASA Image and Video Library

1999-10-21

This is an artist's rendition of an antimatter propulsion system. Matter - antimatter arnihilation offers the highest possible physical energy density of any known reaction substance. It is about 10 billion times more powerful than that of chemical engergy such as hydrogen and oxygen combustion. Antimatter would be the perfect rocket fuel, but the problem is that the basic component of antimatter, antiprotons, doesn't exist in nature and has to manufactured. The process of antimatter development is on-going and making some strides, but production of this as a propulsion system is far into the future.

Artist's concept of Antimatter propulsion system

NASA Technical Reports Server (NTRS)

1999-01-01

This is an artist's rendition of an antimatter propulsion system. Matter - antimatter arnihilation offers the highest possible physical energy density of any known reaction substance. It is about 10 billion times more powerful than that of chemical engergy such as hydrogen and oxygen combustion. Antimatter would be the perfect rocket fuel, but the problem is that the basic component of antimatter, antiprotons, doesn't exist in nature and has to manufactured. The process of antimatter development is on-going and making some strides, but production of this as a propulsion system is far into the future.

Explicit Pore Pressure Material Model in Carbon-Cloth Phenolic

NASA Technical Reports Server (NTRS)

Gutierrez-Lemini, Danton; Ehle, Curt

2003-01-01

An explicit material model that uses predicted pressure in the pores of a carbon-cloth phenolic (CCP) composite has been developed. This model is intended to be used within a finite-element model to predict phenomena specific to CCP components of solid-fuel-rocket nozzles subjected to high operating temperatures and to mechanical stresses that can be great enough to cause structural failures. Phenomena that can be predicted with the help of this model include failures of specimens in restrained-thermal-growth (RTG) tests, pocketing erosion, and ply lifting

Comprehensive modeling of a liquid rocket combustion chamber

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

Nuclear Thermal Propulsion: A Joint NASA/DOE/DOD Workshop

NASA Technical Reports Server (NTRS)

Clark, John S. (Editor)

1991-01-01

Papers presented at the joint NASA/DOE/DOD workshop on nuclear thermal propulsion are compiled. The following subject areas are covered: nuclear thermal propulsion programs; Rover/NERVA and NERVA systems; Low Pressure Nuclear Thermal Rocket (LPNTR); particle bed reactor nuclear rocket; hybrid propulsion systems; wire core reactor; pellet bed reactor; foil reactor; Droplet Core Nuclear Rocket (DCNR); open cycle gas core nuclear rockets; vapor core propulsion reactors; nuclear light bulb; Nuclear rocket using Indigenous Martian Fuel (NIMF); mission analysis; propulsion and reactor technology; development plans; and safety issues.

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.

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

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

Droplet-turbulence interactions in subcritical and supercritical evaporating sprays

NASA Technical Reports Server (NTRS)

Santavicca, Domenic A.; Coy, Edward; Greenfield, Stuart; Song, Young-Hoon

1991-01-01

The objective of this research is to obtain an improved understanding of droplet turbulence interactions in vaporizing liquid sprays under conditions typical of those encountered in liquid fueled rocket engines. The interaction between liquid droplets and the surrounding turbulent gas flow affects droplet dispersion, droplet collisions, droplet vaporization and gas-phase, fuel-oxidant mixing, and therefore has a significant effect on the engine's combustion characteristics. An example of this is the role which droplet-turbulence interactions are believed to play in combustion instabilities. Despite their importance, droplet-turbulence interactions and their effect on liquid fueled rocket engine performance are not well understood. This is particularly true under supercritical conditions, where many conventional concepts, such as surface tension, no longer apply. Our limited understanding of droplet-turbulence interactions, under both subcritical conditions, represents a major limitation in our ability to design improved liquid previously unavailable information and valuable new insights which will directly impact the design of future liquid fueled rocket engines, as well as, allow for the development of significantly improved spray combustion models, making such models useful design tools.

Dr. Robert Goddard

NASA Image and Video Library

2010-01-04

The Goddard Space Flight Center was named in honor of Dr. Robert Goddard, a pioneer in rocket development. Dr. Goddard received patents for a multi-stage rocket and liquid propellants in 1914 and published a paper describing how to reach extreme altitudes six years later. That paper, "A Method of Reaching Extreme Altitudes," detailed methods for raising weather-recording instruments higher than what could be achieved by balloons and explained the mathematical theories of rocket propulsion. The paper, which was published by the Smithsonian Institution, also discussed the possibility of a rocket reaching the moon-a position for which the press ridiculed Goddard. Yet several copies of the report found their way to Europe, and by1927, the German Rocket Society was established, and the German Army began its rocket program in 1931. Goddard, meanwhile, continued his work. By 1926, he had constructed and tested the first rocket using liquid fuel. Goddard's work largely anticipated in technical detail the later German V-2 missiles, including gyroscopic control, steering by means of vanes in the jet stream of the rocket motor, gimbal-steering, power-driven fuel pumps and other devices. NASA Goddard Space Flight Center enables NASA’s mission through four scientific endeavors: Earth Science, Heliophysics, Solar System Exploration, and Astrophysics. Goddard plays a leading role in NASA’s accomplishments by contributing compelling scientific knowledge to advance the Agency’s mission. Follow us on Twitter Join us on Facebook

Heterogeneous fuel for hybrid rocket

NASA Technical Reports Server (NTRS)

Stickler, David B. (Inventor)

1996-01-01

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

Fiberoptic characteristics for extreme operating environments

NASA Technical Reports Server (NTRS)

Delcher, R. C.

1992-01-01

Fiberoptics could offer several major benefits for cryogenic liquid-fueled rocket engines, including lightning immunity, weight reduction, and the possibility of implementing a number of new measurements for engine condition monitoring. The technical feasibility of using fiberoptics in the severe environments posed by cryogenic liquid-fueled rocket engines was determined. The issues of importance and subsequent requirements for this use of fiberoptics were compiled. These included temperature ranges, moisture embrittlement succeptability, and the ability to withstand extreme shock and vibration levels. Different types of optical fibers were evaluated and several types of optical fibers' ability to withstand use in cryogenic liquid-fueled rocket engines was demonstrated through environmental testing of samples. This testing included: cold-bend testing, moisture embrittlement testing, temperature cycling, temperature extremes testing, vibration testing, and shock testing. Three of five fiber samples withstood the tests to a level proving feasibility, and two of these remained intact in all six of the tests. A fiberoptic bundle was also tested, and completed testing without breakage. Preliminary cabling and harnessing for fiber protection was also demonstrated. According to cable manufacturers, the successful -300 F cold bend, vibration, and shock tests are the first instance of any major fiberoptic cable testing below roughly -55 F. This program has demonstrated the basic technical feasibility of implementing optical fibers on cryogenic liquid-fueled rocket engines, and a development plan is included highlighting requirements and issues for such an implementation.

43. HISTORIC VIEW LOOKING SOUTHWEST AT THE TEST STAND WITH ...

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

43. HISTORIC VIEW LOOKING SOUTHWEST AT THE TEST STAND WITH A REDSTONE ROCKET BEING FUELED AND PREPARED FOR TESTING. - Marshall Space Flight Center, Redstone Rocket (Missile) Test Stand, Dodd Road, Huntsville, Madison County, AL

Performance prediction of a ducted rocket combustor

NASA Astrophysics Data System (ADS)

Stowe, Robert

2001-07-01

The ducted rocket is a supersonic flight propulsion system that takes the exhaust from a solid fuel gas generator, mixes it with air, and burns it to produce thrust. To develop such systems, the use of numerical models based on Computational Fluid Dynamics (CFD) is increasingly popular, but their application to reacting flow requires specific attention and validation. Through a careful examination of the governing equations and experimental measurements, a CFD-based method was developed to predict the performance of a ducted rocket combustor. It uses an equilibrium-chemistry Probability Density Function (PDF) combustion model, with a gaseous and a separate stream of 75 nm diameter carbon spheres to represent the fuel. After extensive validation with water tunnel and direct-connect combustion experiments over a wide range of geometries and test conditions, this CFD-based method was able to predict, within a good degree of accuracy, the combustion efficiency of a ducted rocket combustor.

A performance comparison of two small rocket nozzles

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

Chance Encounter with a Stratospheric Kerosene Rocket Plume From Russia Over California

NASA Technical Reports Server (NTRS)

Newman, P. A.; Wilson, J. C.; Ross, M. N.; Brock, C. A.; Sheridan, P. J.; Schoeberl, M. R.; Lait, L. R.; Bui, T. P.; Loewenstein, M.; Podolske, J. R.;

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

Fabry-Perot interferometer development for rocket engine plume spectroscopy

NASA Astrophysics Data System (ADS)

Bickford, R. L.; Madzsar, G.

1990-07-01

This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

Fabry-Perot interferometer development for rocket engine plume spectroscopy

NASA Technical Reports Server (NTRS)

Bickford, R. L.; Madzsar, G.

1990-01-01

This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

A simplified computational fluid-dynamic approach to the oxidizer injector design in hybrid rockets

NASA Astrophysics Data System (ADS)

Di Martino, Giuseppe D.; Malgieri, Paolo; Carmicino, Carmine; Savino, Raffaele

2016-12-01

Fuel regression rate in hybrid rockets is non-negligibly affected by the oxidizer injection pattern. In this paper a simplified computational approach developed in an attempt to optimize the oxidizer injector design is discussed. Numerical simulations of the thermo-fluid-dynamic field in a hybrid rocket are carried out, with a commercial solver, to investigate into several injection configurations with the aim of increasing the fuel regression rate and minimizing the consumption unevenness, but still favoring the establishment of flow recirculation at the motor head end, which is generated with an axial nozzle injector and has been demonstrated to promote combustion stability, and both larger efficiency and regression rate. All the computations have been performed on the configuration of a lab-scale hybrid rocket motor available at the propulsion laboratory of the University of Naples with typical operating conditions. After a preliminary comparison between the two baseline limiting cases of an axial subsonic nozzle injector and a uniform injection through the prechamber, a parametric analysis has been carried out by varying the oxidizer jet flow divergence angle, as well as the grain port diameter and the oxidizer mass flux to study the effect of the flow divergence on heat transfer distribution over the fuel surface. Some experimental firing test data are presented, and, under the hypothesis that fuel regression rate and surface heat flux are proportional, the measured fuel consumption axial profiles are compared with the predicted surface heat flux showing fairly good agreement, which allowed validating the employed design approach. Finally an optimized injector design is proposed.

Computational Analysis of the Combustion Processes in an Axisymmetric, RBCC Flowpath

NASA Technical Reports Server (NTRS)

Steffen, Christopher J., Jr.; Yungster, Shaye

2001-01-01

Computational fluid dynamic simulations have been used to study the combustion processes within an axisymmetric, RBCC flowpath. Two distinct operating modes have been analyzed to date, including the independent ramjet stream (IRS) cycle and the supersonic combustion ramjet (scramJet) cycle. The IRS cycle investigation examined the influence of fuel-air ratio, fuel distribution, and rocket chamber pressure upon the combustion physics and thermal choke characteristics. Results indicate that adjustment of the amount and radial distribution of fuel can control the thermal choke point. The secondary massflow rate was very sensitive to the fuel-air ratio and the rocket chamber pressure. The scramjet investigation examined the influence of fuel-air ratio and fuel injection schedule upon combustion performance estimates. An analysis of the mesh-dependence of these calculations was presented. Jet penetration data was extracted from the three-dimensional simulations and compared favorably with experimental correlations of similar flows. Results indicate that combustion efficiency was very sensitive to the fuel schedule.

Carbide fuels for nuclear thermal propulsion

NASA Astrophysics Data System (ADS)

Matthews, R. B.; Blair, H. T.; Chidester, K. M.; Davidson, K. V.; Stark, W. E.; Storms, E. K.

1991-09-01

A renewed interest in manned exploration of space has revitalized interest in the potential for advancing nuclear rocket technology developed during the 1960's. Carbide fuel performance, melting point, stability, fabricability and compatibility are key technology issues for advanced Nuclear Thermal Propulsion reactors. The Rover fuels development ended with proven carbide fuel forms with demonstrated operating temperatures up to 2700 K for over 100 minutes. The next generation of nuclear rockets will start where the Rover technology ended, but with a more rigorous set of operating requirements including operating lifetime to 10 hours, operating temperatures greater that 3000 K, low fission product release, and compatibility. A brief overview of Rover/NERVA carbide fuel development is presented. A new fuel form with the highest potential combination of operating temperature and lifetime is proposed that consists of a coated uranium carbide fuel sphere with built-in porosity to contain fission products. The particles are dispersed in a fiber reinforced ZrC matrix to increase thermal shock resistance.

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

DTIC Science & Technology

2007-08-01

been demonstrated that a hybrid rocket system using 85% hydrogen peroxide ( HTP ) as the oxidizer and polyethylene as the solid fuel can serve as a cost...As with the tests at Surrey, they used a catalyst pack to decompose the HTP for the ignition. This type of process provides a self-ignition behavior...low regression rate as HTP and polyethylene, so it is difficult to obtain high thrust levels. MARS has the distinction of launching the first

Advanced Chemical Propulsion

NASA Technical Reports Server (NTRS)

Bai, S. Don

2000-01-01

Design, propellant selection, and launch assistance for advanced chemical propulsion system is discussed. Topics discussed include: rocket design, advance fuel and high energy density materials, launch assist, and criteria for fuel selection.

SolSTUS: Solar Source Thermal Upper Stage

NASA Astrophysics Data System (ADS)

This paper was written by members of the Utah State University (USU) Space Systems Design class, fall quarter 1993. The class is funded by NASA and administered by the University Space Research Association (USRA). The focus of the class is to give students some experience in design of space systems and as a source of original ideas for NASA. This paper is a summary of the work done by members of the Space Systems Design class during the opening phase of the course. The class was divided into groups to work on different areas of the Solar Thermal Rocket (STR) booster in order to produce a design reference mission that would identify the key design issues. The design reference mission focused upon a small satellite mission to Mars. There are several critical components in a Solar Thermal Rocket. STR's produce a very low thrust, but have a high specific impulse, meaning that they take longer to reach the desired orbit, but use a lot less fuel in doing it. The complexity of the rocket is discussed in this paper. Some of the more critical design problems discussed are: (1) the structural and optical complexity of collecting and focusing sunlight onto a specific point, (2) long term storage of fuel (liquid hydrogen), (3) attitude control while thrusting in an elliptical orbit and orienting the mirrors to collect sunlight, and (4) power and communications for the rocket and it's internal systems. The design reference mission discussed here is a very general mission to Mars. A first order trajectory design has been done and a possible basic science payload for Mars has been suggested. This paper summarizes the design reference mission (DRM) formulated by the USU students during fall quarter and identifies major design challenges that will confront the design team during the next two quarters here at USU.

SolSTUS: Solar Source Thermal Upper Stage

NASA Technical Reports Server (NTRS)

1994-01-01

This paper was written by members of the Utah State University (USU) Space Systems Design class, fall quarter 1993. The class is funded by NASA and administered by the University Space Research Association (USRA). The focus of the class is to give students some experience in design of space systems and as a source of original ideas for NASA. This paper is a summary of the work done by members of the Space Systems Design class during the opening phase of the course. The class was divided into groups to work on different areas of the Solar Thermal Rocket (STR) booster in order to produce a design reference mission that would identify the key design issues. The design reference mission focused upon a small satellite mission to Mars. There are several critical components in a Solar Thermal Rocket. STR's produce a very low thrust, but have a high specific impulse, meaning that they take longer to reach the desired orbit, but use a lot less fuel in doing it. The complexity of the rocket is discussed in this paper. Some of the more critical design problems discussed are: (1) the structural and optical complexity of collecting and focusing sunlight onto a specific point, (2) long term storage of fuel (liquid hydrogen), (3) attitude control while thrusting in an elliptical orbit and orienting the mirrors to collect sunlight, and (4) power and communications for the rocket and it's internal systems. The design reference mission discussed here is a very general mission to Mars. A first order trajectory design has been done and a possible basic science payload for Mars has been suggested. This paper summarizes the design reference mission (DRM) formulated by the USU students during fall quarter and identifies major design challenges that will confront the design team during the next two quarters here at USU.

Hybrid Propulsion Technology Program

NASA Technical Reports Server (NTRS)

Jensen, G. E.; Holzman, A. L.

1990-01-01

Future launch systems of the United States will require improvements in booster safety, reliability, and cost. In order to increase payload capabilities, performance improvements are also desirable. The hybrid rocket motor (HRM) offers the potential for improvements in all of these areas. The designs are presented for two sizes of hybrid boosters, a large 4.57 m (180 in.) diameter booster duplicating the Advanced Solid Rocket Motor (ASRM) vacuum thrust-time profile and smaller 2.44 m (96 in.), one-quater thrust level booster. The large booster would be used in tandem, while eight small boosters would be used to achieve the same total thrust. These preliminary designs were generated as part of the NASA Hybrid Propulsion Technology Program. This program is the first phase of an eventual three-phaes program culminating in the demonstration of a large subscale engine. The initial trade and sizing studies resulted in preferred motor diameters, operating pressures, nozzle geometry, and fuel grain systems for both the large and small boosters. The data were then used for specific performance predictions in terms of payload and the definition and selection of the requirements for the major components: the oxidizer feed system, nozzle, and thrust vector system. All of the parametric studies were performed using realistic fuel regression models based upon specific experimental data.

Optical Phenomena Observed upon Some Launches of Russian Rockets

NASA Astrophysics Data System (ADS)

Kozlov, S. I.; Nilolaishvili, S. Sh.; Platov, Yu. V.

2018-01-01

In this paper, unusual optical phenomena observed in our country and abroad upon launches of Russian rockets are discussed and interpreted: they are regarded as the aftereffects of sunlight scattering by gas-dust clouds created by rocket fuel combustion products in different modes of engine operation. The results of instrumental observations of the clouds can be used to study physical processes in the upper atmosphere.

Cooled Ceramic Composite Panel Tested Successfully in Rocket Combustion Facility

NASA Technical Reports Server (NTRS)

Jaskowiak, Martha H.

2003-01-01

Regeneratively cooled ceramic matrix composite (CMC) structures are being considered for use along the walls of the hot-flow paths of rocket-based or turbine-based combined-cycle propulsion systems. They offer the combined benefits of substantial weight savings, higher operating temperatures, and reduced coolant requirements in comparison to components designed with traditional metals. These cooled structures, which use the fuel as the coolant, require materials that can survive aggressive thermal, mechanical, acoustic, and aerodynamic loads while acting as heat exchangers, which can improve the efficiency of the engine. A team effort between the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and various industrial partners has led to the design, development, and fabrication of several types of regeneratively cooled panels. The concepts for these panels range from ultra-lightweight designs that rely only on CMC tubes for coolant containment to more maintainable designs that incorporate metal coolant containment tubes to allow for the rapid assembly or disassembly of the heat exchanger. One of the cooled panels based on an all-CMC design was successfully tested in the rocket combustion facility at Glenn. Testing of the remaining four panels is underway.

Historical problem areas lessons learned

NASA Technical Reports Server (NTRS)

Sackheim, Bob; Fester, Dale A.

1991-01-01

Historical problem areas in space transportation propulsion technology are identified in viewgraph form. Problem areas discussed include materials compatibility, contamination, pneumatic/feed system flow instabilities, instabilities in rocket engine combustion and fuel sloshing, exhaust plume interference, composite rocket nozzle failure, and freeze/thaw damage.

Closed-loop thrust and pressure profile throttling of a nitrous oxide/hydroxyl-terminated polybutadiene hybrid rocket motor

NASA Astrophysics Data System (ADS)

Peterson, Zachary W.

Hybrid motors that employ non-toxic, non-explosive components with a liquid oxidizer and a solid hydrocarbon fuel grain have inherently safe operating characteristics. The inherent safety of hybrid rocket motors offers the potential to greatly reduce overall operating costs. Another key advantage of hybrid rocket motors is the potential for in-flight shutdown, restart, and throttle by controlling the pressure drop between the oxidizer tank and the injector. This research designed, developed, and ground tested a closed-loop throttle controller for a hybrid rocket motor using nitrous oxide and hydroxyl-terminated polybutadiene as propellants. The research simultaneously developed closed-loop throttle algorithms and lab scale motor hardware to evaluate the fidelity of the throttle simulations and algorithms. Initial open-loop motor tests were performed to better classify system parameters and to validate motor performance values. Deep-throttle open-loop tests evaluated limits of stable thrust that can be achieved on the test hardware. Open-loop tests demonstrated the ability to throttle the motor to less than 10% of maximum thrust with little reduction in effective specific impulse and acoustical stability. Following the open-loop development, closed-loop, hardware-in-the-loop tests were performed. The closed-loop controller successfully tracked prescribed step and ramp command profiles with a high degree of fidelity. Steady-state accuracy was greatly improved over uncontrolled thrust.

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

NASA Technical Reports Server (NTRS)

Pavli, A. J.

1979-01-01

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

State and prospects of solid propellant rocket development

NASA Astrophysics Data System (ADS)

Kukushkin, V. Kh.

1992-07-01

An overview is presented of aspects of solid-propellant rocket engine (SPRE) development with individual treatment given to sustainer and spacecraft SPRE technologies. The paper focuses on low-modulus fuels of composite solid propellant, requirements for adhesion stability, and enhancement of the power characteristics of solid propellants. R&D activities are described that relate to the use of SPREs with extending nozzles and to the design of ultradimensional nozzles for upper-stage engines. Other developments for the SPREs include engines with separate loading and pasty fuel applications, and progress is reported in the direction of detonation SPREs. The SPREs using pasty propellants provide good control over thrust characteristics and fuel qualities. A device is incorporated that assures fuel burning in the combustion region and reliable ignition during restarting of these engines.

Space reliability technology - A historical perspective

NASA Technical Reports Server (NTRS)

Cohen, H.

1984-01-01

The progressive improvements in reliability of launch vehicles is traced from the Vanguard rocket to the STS. The Vanguard, built with minimal redundancy and a high mass ratio, was used as an operational vehicle midway through its test program in an attempt to meet the perceived challenge represented by the Sputnik. The fourth Vanguard failed due to inadequate contamination prevention and lack of inspection ports. Automatic firing sequences were adopted for the Titan rockets, which were an order of magnitude larger than the Vanguard and therefore had room for interior inspections. Qualification testing and reporting were introduced for components, along with X ray inspection of fuel tank welds. Dual systems were added for flight critical components when the Titan became man-rated for the Gemini program. Designs incorporated full failure mode effects and criticality analyses for the Apollo program, which exposed the limits of applicability of numerical reliability models. Fault tree analyses and program milestone reviews were initiated. The worth of man-in-the-loop in space activities for reliability was demonstrated with the rescue of Skylab after solar panel and meteoroid shield failures. It is now the reliability of the payload, rather than the vehicle, that is questioned for Shuttle launches.

Gas-Generator Augmented Expander Cycle Rocket Engine

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2011-01-01

An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.

Revised Point of Departure Design Options for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Fittje, James E.; Borowski, Stanley K.; Schnitzler, Bruce

2015-01-01

In an effort to further refine potential point of departure nuclear thermal rocket engine designs, four proposed engine designs representing two thrust classes and utilizing two different fuel matrix types are designed and analyzed from both a neutronics and thermodynamic cycle perspective. Two of these nuclear rocket engine designs employ a tungsten and uranium dioxide cermet (ceramic-metal) fuel with a prismatic geometry based on the ANL-200 and the GE-710, while the other two designs utilize uranium-zirconium-carbide in a graphite composite fuel and a prismatic fuel element geometry developed during the Rover/NERVA Programs. Two engines are analyzed for each fuel type, a small criticality limited design and a 111 kN (25 klbf) thrust class engine design, which has been the focus of numerous manned mission studies, including NASA's Design Reference Architecture 5.0. slightly higher T/W ratios, but they required substantially more 235U.

Effect of grain port length-diameter ratio on combustion performance in hybrid rocket motors

NASA Astrophysics Data System (ADS)

Cai, Guobiao; Zhang, Yuanjun; Tian, Hui; Wang, Pengfei; Yu, Nanjia

2016-11-01

The objectives of this study are to develop a more accurate regression rate considering the oxidizer mass flow and the fuel grain geometry configuration with numerical and experimental investigations in polyethylene (PE)/90% hydrogen peroxide (HP) hybrid rocket. Firstly, a 2-D axisymmetric CFD model with turbulence, chemistry reaction, solid-gas coupling is built to investigate the combustion chamber internal flow structure. Then a more accurate regression formula is proposed and the combustion efficiency changing with the length-diameter ratio is studied. A series experiments are conducted in various oxidizer mass flow to analyze combustion performance including the regression rate and combustion efficiency. The regression rates are measured by the fuel mass reducing and diameter changing. A new regression rate formula considering the fuel grain configuration is proposed in this paper. The combustion efficiency increases with the length-diameter ratio changing. To improve the performance of a hybrid rocket motor, the port length-diameter ratio is suggested 10-12 in the paper.

Exploring Mars: The Ares Payload Service (APS)

NASA Astrophysics Data System (ADS)

Bowen, Justin; Lusignan, Bruce

1999-08-01

In last year's Mars Society convention we introduced the results of five years of studies of space launch capability for the second millennium. We concluded that Single Stage to Orbit (SSTO) vehicles such as the Delta Clipper X33, and X34 cannot make it to orbit from the Earth's surface. Whether taking off vertically or horizontally or landing vertically or horizontally, the rocket equations, the performance of available fuels, and the realities of the weight and strength of materials leave no margin for payload. The promised savings from SSTO systems are illusory. However, a configuration that is able to deliver useful payload to orbit is the Single step to Orbit, SsTO, a rocket plane that is released fully fueled, from 35,000 to 40,000 feet altitude. Three approaches have been proposed. The Hot'l and Molnya Corporation designs carry the fueled rocket plane to altitude on the back of a carrier aircraft. In this design the carrier aircraft is Russia's Antonov 225 the world's largest cargo plane. The rocket plane is a modified version of the Buran, Russia's own space shuttle. Another configuration is Kelly Aviation's concept in which the fully fueled rocket plane is towed to altitude by the cargo plane and then released. A third approach is based on the early "X" planes, which were dropped from the belly of the carrier plane. While the rocket equations indicate that these three concepts can deliver useful payloads, the Stanford review found significant advantages to the approach of Pioneer Rocket, in which the rocket plane flies up to the carrier plane with conventional jet engines, docks, and then loads on the oxidizer for the flight to orbit. This architecture has more reasonable abort modes in case of system failure in either aircraft and can deliver a larger final payload to orbit for a given sized carrier. The Stanford recommendation is that the carrier aircraft be the Antonov 225. A design based on this was presented in a report last year. Refinements to the design notably an improved re-entry cooling system and fueling stability analysis were done this year. More technical detail and a proposed international consortium to develop the SSTO is presented in another session of this year's Mars convention. We believe that there will be no human exploration of Mars based on the Shuttle or Expendable launch vehicles, and no resources available except for a cooperative international program. However, just as the world is learning to cooperate in peacekeeping, we hold out the hope that similar cooperation will develop for Mars exploration. With that in mind, this year we asked the question- "How will the human mission get to Mars if it has to use the SsTO for transportation?"

Exploring Mars: the Ares Payload Service (APS)

NASA Technical Reports Server (NTRS)

Bowen, Justin; Lusignan, Bruce

1999-01-01

In last year's Mars Society convention we introduced the results of five years of studies of space launch capability for the second millennium. We concluded that Single Stage to Orbit (SSTO) vehicles such as the Delta Clipper X33, and X34 cannot make it to orbit from the Earth's surface. Whether taking off vertically or horizontally or landing vertically or horizontally, the rocket equations, the performance of available fuels, and the realities of the weight and strength of materials leave no margin for payload. The promised savings from SSTO systems are illusory. However, a configuration that is able to deliver useful payload to orbit is the Single step to Orbit, SsTO, a rocket plane that is released fully fueled, from 35,000 to 40,000 feet altitude. Three approaches have been proposed. The Hot'l and Molnya Corporation designs carry the fueled rocket plane to altitude on the back of a carrier aircraft. In this design the carrier aircraft is Russia's Antonov 225 the world's largest cargo plane. The rocket plane is a modified version of the Buran, Russia's own space shuttle. Another configuration is Kelly Aviation's concept in which the fully fueled rocket plane is towed to altitude by the cargo plane and then released. A third approach is based on the early "X" planes, which were dropped from the belly of the carrier plane. While the rocket equations indicate that these three concepts can deliver useful payloads, the Stanford review found significant advantages to the approach of Pioneer Rocket, in which the rocket plane flies up to the carrier plane with conventional jet engines, docks, and then loads on the oxidizer for the flight to orbit. This architecture has more reasonable abort modes in case of system failure in either aircraft and can deliver a larger final payload to orbit for a given sized carrier. The Stanford recommendation is that the carrier aircraft be the Antonov 225. A design based on this was presented in a report last year. Refinements to the design notably an improved re-entry cooling system and fueling stability analysis were done this year. More technical detail and a proposed international consortium to develop the SSTO is presented in another session of this year's Mars convention. We believe that there will be no human exploration of Mars based on the Shuttle or Expendable launch vehicles, and no resources available except for a cooperative international program. However, just as the world is learning to cooperate in peacekeeping, we hold out the hope that similar cooperation will develop for Mars exploration. With that in mind, this year we asked the question- "How will the human mission get to Mars if it has to use the SsTO for transportation?"

A real-time guidance algorithm for aerospace plane optimal ascent to low earth orbit

NASA Technical Reports Server (NTRS)

Calise, A. J.; Flandro, G. A.; Corban, J. E.

1989-01-01

Problems of onboard trajectory optimization and synthesis of suitable guidance laws for ascent to low Earth orbit of an air-breathing, single-stage-to-orbit vehicle are addressed. A multimode propulsion system is assumed which incorporates turbojet, ramjet, Scramjet, and rocket engines. An algorithm for generating fuel-optimal climb profiles is presented. This algorithm results from the application of the minimum principle to a low-order dynamic model that includes angle-of-attack effects and the normal component of thrust. Maximum dynamic pressure and maximum aerodynamic heating rate constraints are considered. Switching conditions are derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another. A nonlinear transformation technique is employed to derived a feedback controller for tracking the computed trajectory. Numerical results illustrate the nature of the resulting fuel-optimal climb paths.

Antimatter Propulsion Developed by NASA

NASA Technical Reports Server (NTRS)

1999-01-01

This Quick Time movie shows possible forms of an antimatter propulsion system being developed by NASA. Antimatter annihilation offers the highest possible physical energy density of any known reaction substance. It is about 10 billion times more powerful than that of chemical energy such as hydrogen and oxygen combustion. Antimatter would be the perfect rocket fuel, but the problem is that the basic component of antimatter, antiprotons, doesn't exist in nature and has to manufactured. The process of antimatter development is ongoing and making some strides, but production of this as a propulsion system is far into the future.

Design of a Six Degree of Freedom Thrust Sensor for a Hybrid Rocket

NASA Astrophysics Data System (ADS)

McGehee, Tripp

2005-03-01

A hybrid rocket is composed of a solid fuel and a separate liquid or gaseous oxidizer. These rockets may be throttled like liquid rockets, are safer than solid rockets, and are much less complex than liquid rockets. However, hybrid rockets produce thrust oscillations that are not practical for large scale use. A lab scale hybrid rocket at the University of Arkansas at Little Rock (UALR) Hybrid Rocket Facility is used to develop sensors to measure physical properties of hybrid rockets. Research is currently being conducted to design a six degree of freedom force sensor to measure the thrust and torque in all three spatial dimensions. The current design mounts the rocket in a rigid cage and connects the cage to a solid table by six sensor legs. The legs utilize strain gauges and a Wheatstone bridge to produce a voltage proportional to the force on the leg. A detailed description of the cage design and the design process will be given.

A Versatile Rocket Engine Hot Gas Facility

NASA Technical Reports Server (NTRS)

Green, James M.

1993-01-01

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

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

NASA Astrophysics Data System (ADS)

Zhang, Shuai; Hu, Fan; Zhang, Weihua

2016-10-01

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

Nonlinear Control of a Reusable Rocket Engine for Life Extension

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.; Holmes, Michael S.; Ray, Asok

1998-01-01

This paper presents the conceptual development of a life-extending control system where the objective is to achieve high performance and structural durability of the plant. A life-extending controller is designed for a reusable rocket engine via damage mitigation in both the fuel (H2) and oxidizer (O2) turbines while achieving high performance for transient responses of the combustion chamber pressure and the O2/H2 mixture ratio. The design procedure makes use of a combination of linear and nonlinear controller synthesis techniques and also allows adaptation of the life-extending controller module to augment a conventional performance controller of the rocket engine. The nonlinear aspect of the design is achieved using non-linear parameter optimization of a prescribed control structure. Fatigue damage in fuel and oxidizer turbine blades is primarily caused by stress cycling during start-up, shutdown, and transient operations of a rocket engine. Fatigue damage in the turbine blades is one of the most serious causes for engine failure.

High-temperature, high-pressure optical port for rocket engine applications

NASA Technical Reports Server (NTRS)

Delcher, Ray; Nemeth, ED; Powers, W. T.

1993-01-01

This paper discusses the design, fabrication, and test of a window assembly for instrumentation of liquid-fueled rocket engine hot gas systems. The window was designed to allow optical measurements of hot gas in the SSME fuel preburner and appears to be the first window designed for application in a rocket engine hot gas system. Such a window could allow the use of a number of remote optical measurement technologies including: Raman temperature and species concentration measurement, Raleigh temperature measurements, flame emission monitoring, flow mapping, laser-induced florescence, and hardware imaging during engine operation. The window assembly has been successfully tested to 8,000 psi at 1000 F and over 11,000 psi at room temperature. A computer stress analysis shows the window will withstand high temperature and cryogenic thermal shock.

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

NASA Technical Reports Server (NTRS)

Pavli, A. J.

1979-01-01

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

Laser diagnostics for NTP fuel corrosion studies

NASA Technical Reports Server (NTRS)

Wantuck, Paul J.; Butt, D. P.; Sappey, A. D.

1993-01-01

Viewgraphs and explanations on laser diagnostics for nuclear thermal propulsion (NTP) fuel corrosion studies are presented. Topics covered include: NTP fuels; U-Zr-C system corrosion products; planar laser-induced fluorescence (PLIF); utilization of PLIF for corrosion product characterization of nuclear thermal rocket fuel elements under test; ZrC emission spectrum; and PLIF imaging of ZrC plume.

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

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.

2017-01-01

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

Combustion Performance of a Staged Hybrid Rocket with Boron addition

NASA Astrophysics Data System (ADS)

Lee, D.; Lee, C.

2018-04-01

In this paper, the effect of boron on overall system specific impulse was investigated. Additionally, a series of combustion tests was carried out to analyze and evaluate the effect of boron addition on O/F variation and radial temperature profiles. To maintain the hybrid rocket engine advantages, upper limit of boron contents in solid fuel was set to be 10 wt%. The results also suggested that, when adding boron to solid fuel, it helped to provide more uniform radial temperature distribution and also to increase specific impulse by 3.2%.

SPOC Benchmark Case: SNRE Model

DOE Office of Scientific and Technical Information (OSTI.GOV)

Vishal Patel; Michael Eades; Claude Russel Joyner II

The Small Nuclear Rocket Engine (SNRE) was modeled in the Center for Space Nuclear Research’s (CSNR) Space Propulsion Optimization Code (SPOC). SPOC aims to create nuclear thermal propulsion (NTP) geometries quickly to perform parametric studies on design spaces of historic and new NTP designs. The SNRE geometry was modeled in SPOC and a critical core with a reasonable amount of criticality margin was found. The fuel, tie-tubes, reflector, and control drum masses were predicted rather well. These are all very important for neutronics calculations so the active reactor geometries created with SPOC can continue to be trusted. Thermal calculations ofmore » the average and hot fuel channels agreed very well. The specific impulse calculations used historically and in SPOC disagree so mass flow rates and impulses differed. Modeling peripheral and power balance components that do not affect nuclear characteristics of the core is not a feature of SPOC and as such, these components should continue to be designed using other tools. A full paper detailing the available SNRE data and comparisons with SPOC outputs will be submitted as a follow-up to this abstract.« less

Experimental investigation of paraffin-based fuels for hybrid rocket propulsion

NASA Astrophysics Data System (ADS)

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

2013-03-01

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

Moving Material into Space Without Rockets.

ERIC Educational Resources Information Center

Cheng, R. S.; Trefil, J. S.

1985-01-01

In response to conventional rocket demands on fuel supplies, electromagnetic launches were developed to give payloads high velocity using a stationary energy source. Several orbital mechanics problems are solved including a simple problem (radial launch with no rotation) and a complex problem involving air resistance and gravity. (DH)

Chance Encounter with a Stratospheric Kerosene Rocket Plume from Russia over California

NASA Technical Reports Server (NTRS)

Newman, P. A.; Wilson, J. C.; Ross, M. N.; Brock, C.; Sheridan, P.; Schoeberl, M. R.; Lait, L. R.; Bui, T. P.; Loewenstein, M.

1999-01-01

During a routine ER-2 aircraft high-altitude test flight on April 18, 1997, an unusual aerosol cloud was detected at 20 km altitude near the California coast at about 370 degrees N latitude. Not visually observed by the ER-2 pilot, the cloud was characterized bv high concentration of soot and sulfate aerosol in a region over 100 km in horizontal extent indicating that the source of the plume was a large hydrocarbon fueled vehicle, most likely a launch vehicle powered only by rocket motors burning liquid oxygen and kerosene. Two Russian Soyuz rockets could conceivably have produced the plume. The first was launched from the Baikonur Cosmodrome, Kazakhstan on April 6th; the second was launched from Plesetsk, Russia on April 9. Air parcel trajectory calculations and long-lived tracer gas concentrations in the cloud indicate that the Baikonur rocket launch is the most probable source of the plume. The parcel trajectory calculations do not unambiguously trace the transport of the Soyuz plume from Asia to North America, illustrating serious flaws in the point-to-point trajectory calculations. This chance encounter represents the only measurement of the stratospheric effects of emissions from a rocket powered exclusively with hydrocarbon fuel.

High energy-density liquid rocket fuel performance

NASA Technical Reports Server (NTRS)

Rapp, Douglas C.

1990-01-01

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

Design and Fabrication of the ISTAR Direct-Connect Combustor Experiment at the NASA Hypersonic Tunnel Facility

NASA Technical Reports Server (NTRS)

Lee, Jin-Ho; Krivanek, Thomas M.

2005-01-01

The Integrated Systems Test of an Airbreathing Rocket (ISTAR) project was a flight demonstration project initiated to advance the state of the art in Rocket Based Combined Cycle (RBCC) propulsion development. The primary objective of the ISTAR project was to develop a reusable air breathing vehicle and enabling technologies. This concept incorporated a RBCC propulsion system to enable the vehicle to be air dropped at Mach 0.7 and accelerated up to Mach 7 flight culminating in a demonstration of hydrocarbon scramjet operation. A series of component experiments was planned to reduce the level of risk and to advance the technology base. This paper summarizes the status of a full scale direct connect combustor experiment with heated endothermic hydrocarbon fuels. This is the first use of the NASA GRC Hypersonic Tunnel facility to support a direct-connect test. The technical and mechanical challenges involved with adapting this facility, previously used only in the free-jet configuration, for use in direct connect mode will be also described.

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

NASA Technical Reports Server (NTRS)

Jones, Ronald E.

2012-01-01

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

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

NASA Technical Reports Server (NTRS)

Jones, Ronald E.; Carter, Robert W.

2012-01-01

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

Ley, Willy (1906-69)

NASA Astrophysics Data System (ADS)

Murdin, P.

2000-11-01

Rocket scientist, writer, born in Berlin, Germany. Inspired by reading a work by the space pioneer, HERMANN OBERTH, Ley founded the German Society for Space Travel (1927), enrolled WERNHER VON BRAUN, and helped develop the liquid-fuel rocket. Fled to the USA, and became a science writer, including science fiction and film scripts....

Technicians Manufacture a Nozzle for the Kiwi B-1-B Engine

NASA Image and Video Library

1964-05-21

Technicians manufacture a nozzle for the Kiwi B-1-B nuclear rocket engine in the Fabrication Shop’s vacuum oven at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Nuclear Engine for Rocket Vehicle Applications (NERVA) was a joint NASA and Atomic Energy Commission (AEC) endeavor to develop a nuclear-powered rocket for both long-range missions to Mars and as a possible upper-stage for the Apollo Program. The early portion of the program consisted of basic reactor and fuel system research. This was followed by a series of Kiwi reactors built to test basic nuclear rocket principles in a non-flying nuclear engine. The next phase, NERVA, would create an entire flyable engine. The final phase of the program, called Reactor-In-Flight-Test, would be an actual launch test. The AEC was responsible for designing the nuclear reactor and overall engine. NASA Lewis was responsible for developing the liquid-hydrogen fuel system. The turbopump, which pumped the fuels from the storage tanks to the engine, was the primary tool for restarting the engine. The NERVA had to be able to restart in space on its own using a safe preprogrammed startup system. Lewis researchers endeavored to design and test this system. This non-nuclear Kiwi engine, seen here, was being prepared for tests at Lewis’ High Energy Rocket Engine Research Facility (B-1) located at Plum Brook Station. The tests were designed to start an unfueled Kiwi B-1-B reactor and its Aerojet Mark IX turbopump without any external power.

Schlieren and OH* chemiluminescence imaging of combustion in a turbulent boundary layer over a solid fuel

NASA Astrophysics Data System (ADS)

Jens, Elizabeth T.; Miller, Victor A.; Cantwell, Brian J.

2016-03-01

Combustion in a turbulent boundary layer over a solid fuel is studied using simultaneous schlieren and OH* chemiluminescence imaging. The flow configuration is representative of a hybrid rocket motor combustor. Six different hydrocarbon fuels, including both classical hybrid rocket fuels and a high regression rate fuel (paraffin wax), are burned in an undiluted oxygen free-stream at pressures ranging from atmospheric to 1524.2 kPa (221.1 psi). A detailed explanation of methods for registering the schlieren and OH* chemiluminescence images to one another is presented, and additionally, details of the routines used to extract flow features of interest (like the boundary layer height and flame location) are provided. At atmospheric pressure, the boundary layer location is consistent between all fuels; however, the flame location varies for each fuel. The flame zone appears to be smoothly distributed over the fuel surface at atmospheric pressure. At elevated pressures and correspondingly increased Dahmköhler number (but at constant Reynolds number), flame morphology is markedly different, exhibiting large rollers in a shear layer above the fuel grain and finer structures in the flame. The chemiluminescence intensity is found to be roughly proportional to the fuel burn rate at both atmospheric and elevated chamber pressures.

Experimental Study of Ballistic-Missile Base Heating with Operating Rocket

NASA Technical Reports Server (NTRS)

Nettle, J. Cary

1958-01-01

A rocket of the 1000-pound-thrust class using liquid oxygen and JP-4 fuel as propellant was installed in the Lewis 8- by 6-foot tunnel to permit a controlled study of some of the factors affecting the heating of a rocket-missile base. Temperatures measured in the base region are presented from findings of three motor extension lengths relative to the base. Data are also presented for two combustion efficiency levels in the rocket motor. Temperature as high as 1200 F was measured in the base region because of the ignition of burnable rocket gases. combustibles that are dumped into the base by accessories seriously aggravate the base-burning temperature rise.

The use of programmable logic controllers (PLC) for rocket engine component testing

NASA Technical Reports Server (NTRS)

Nail, William; Scheuermann, Patrick; Witcher, Kern

1991-01-01

Application of PLCs to the rocket engine component testing at a new Stennis Space Center Component Test Facility is suggested as an alternative to dedicated specialized computers. The PLC systems are characterized by rugged design, intuitive software, fault tolerance, flexibility, multiple end device options, networking capability, and built-in diagnostics. A distributed PLC-based system is projected to be used for testing LH2/LOx turbopumps required for the ALS/NLS rocket engines.

Physics and potentials of fissioning plasmas for space power and propulsion

NASA Technical Reports Server (NTRS)

Thom, K.; Schwenk, F. C.; Schneider, R. T.

1976-01-01

Fissioning uranium plasmas are the nuclear fuel in conceptual high-temperature gaseous-core reactors for advanced rocket propulsion in space. A gaseous-core nuclear rocket would be a thermal reactor in which an enriched uranium plasma at about 10,000 K is confined in a reflector-moderator cavity where it is nuclear critical and transfers its fission power to a confining propellant flow for the production of thrust at a specific impulse up to 5000 sec. With a thrust-to-engine weight ratio approaching unity, the gaseous-core nuclear rocket could provide for propulsion capabilities needed for manned missions to the nearby planets and for economical cislunar ferry services. Fueled with enriched uranium hexafluoride and operated at temperatures lower than needed for propulsion, the gaseous-core reactor scheme also offers significant benefits in applications for space and terrestrial power. They include high-efficiency power generation at low specific mass, the burnup of certain fission products and actinides, the breeding of U-233 from thorium with short doubling times, and improved convenience of fuel handling and processing in the gaseous phase.

Thermal analysis of regenerative-cooled pylon in multi-mode rocket based combined cycle engine

NASA Astrophysics Data System (ADS)

Yan, Dekun; He, Guoqiang; Li, Wenqiang; Zhang, Duo; Qin, Fei

2018-07-01

Combining pylon injector with rocket is an effective method to achieve efficient mixing and combustion in the RBCC engine. This study designs a fuel pylon with active cooling structure, and numerically investigates the coupled heat transfer between active cooling process in the pylon and combustion in the combustor in different modes. Effect of the chemical reaction of the fuel on the flow, heat transfer and physical characteristics is also discussed. The numerical results present a good agreement with the experimental data. Results indicate that drastic supplementary combustion caused by rocket gas and secondary combustion caused by the fuel injection from the pylon result in severe thermal load on the pylon. Although regenerative cooling without cracking can reduce pylon's temperature below the allowable limit, a high-temperature area appears in the middle and nail section of the pylon due to the coolant's insufficient convective heat transfer coefficient. Comparatively, endothermic cracking can provide extra chemical heat sink for the coolant and low velocity contributes to prolong the reaction time to increase the heat absorption from chemical reaction, which further lowers and unifies the pylon surface temperature.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

NASA Technical Reports Server (NTRS)

Litchford, Ron J.; Foote, John P.; Clifton, W. B.; Hickman, Robert R.; Wang, Ten-See; Dobson, Christopher C.

2011-01-01

An arc-heater driven hyperthermal convective environments simulator was recently developed and commissioned for long duration hot hydrogen exposure of nuclear thermal rocket materials. This newly established non-nuclear testing capability uses a high-power, multi-gas, wall-stabilized constricted arc-heater to produce hightemperature pressurized hydrogen flows representative of nuclear reactor core environments, excepting radiation effects, and is intended to serve as a low-cost facility for supporting non-nuclear developmental testing of hightemperature fissile fuels and structural materials. The resulting reactor environments simulator represents a valuable addition to the available inventory of non-nuclear test facilities and is uniquely capable of investigating and characterizing candidate fuel/structural materials, improving associated processing/fabrication techniques, and simulating reactor thermal hydraulics. This paper summarizes facility design and engineering development efforts and reports baseline operational characteristics as determined from a series of performance mapping and long duration capability demonstration tests. Potential follow-on developmental strategies are also suggested in view of the technical and policy challenges ahead. Keywords: Nuclear Rocket Engine, Reactor Environments, Non-Nuclear Testing, Fissile Fuel Development.

Spark Ignition of Combustible Vapor in a Plastic Bottle as a Demonstration of Rocket Propulsion

NASA Astrophysics Data System (ADS)

Mattox, J. R.

2017-01-01

I report an innovation that provides a compelling demonstration of rocket propulsion, appropriate for students of physics and other physical sciences. An electrical spark is initiated from a distance to cause the deflagration of a combustible vapor mixed with air in a lightweight plastic bottle that is consequently propelled as a rocket by the release of combustion products, i.e., a "whoosh rocket." My recommendation is that the standard fuel for pedagogical whoosh demonstrations be isopropanol, and the recommended vessel is the 3.8-L high-density polyethylene (HDPE) bottle.

Optimization of the rocket mode trajectory in a rocket based combined cycle (RBCC) engine powered SSTO vehicle

NASA Astrophysics Data System (ADS)

Foster, Richard W.

1989-07-01

The application of rocket-based combined cycle (RBCC) engines to booster-stage propulsion, in combination with all-rocket second stages in orbital-ascent missions, has been studied since the mid-1960s; attention is presently given to the case of the 'ejector scramjet' RBCC configuration's application to SSTO vehicles. While total mass delivered to initial orbit is optimized at Mach 20, payload delivery capability to initial orbit optimizes at Mach 17, primarily due to the reduction of hydrogen fuel tankage structure, insulation, and thermal protection system weights.

Moving, Moving, Moving- A Giant Rocket Fuel Tank

NASA Image and Video Library

2016-10-07

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

Modeling and Simulation of a Nuclear Fuel Element Test Section

NASA Technical Reports Server (NTRS)

Moran, Robert P.; Emrich, William

2011-01-01

"The Nuclear Thermal Rocket Element Environmental Simulator" test section closely simulates the internal operating conditions of a thermal nuclear rocket. The purpose of testing is to determine the ideal fuel rod characteristics for optimum thermal heat transfer to their hydrogen cooling/working fluid while still maintaining fuel rod structural integrity. Working fluid exhaust temperatures of up to 5,000 degrees Fahrenheit can be encountered. The exhaust gas is rendered inert and massively reduced in temperature for analysis using a combination of water cooling channels and cool N2 gas injectors in the H2-N2 mixer portion of the test section. An extensive thermal fluid analysis was performed in support of the engineering design of the H2-N2 mixer in order to determine the maximum "mass flow rate"-"operating temperature" curve of the fuel elements hydrogen exhaust gas based on the test facilities available cooling N2 mass flow rate as the limiting factor.

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

NASA Technical Reports Server (NTRS)

Verma, Satyajit

2006-01-01

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

Wavelength modulation spectroscopy near 5 μm for carbon monoxide sensing in a high-pressure kerosene-fueled liquid rocket combustor

NASA Astrophysics Data System (ADS)

Lee, Daniel D.; Bendana, Fabio A.; Schumaker, S. Alexander; Spearrin, R. Mitchell

2018-05-01

A laser absorption sensor was developed for carbon monoxide (CO) sensing in high-pressure, fuel-rich combustion gases associated with the internal conditions of hydrocarbon-fueled liquid bipropellant rockets. An absorption feature near 4.98 μm, comprised primarily of two rovibrational lines from the P-branch of the fundamental band, was selected to minimize temperature sensitivity and spectral interference with other combustion gas species at the extreme temperatures (> 3000 K) and pressures (> 50 atm) in the combustion chamber environment. A scanned wavelength modulation spectroscopy technique (1 f-normalized 2 f detection) is utilized to infer species concentration from CO absorption, and mitigate the influence of non-absorption transmission losses and noise associated with the harsh sooting combustor environment. To implement the sensing strategy, a continuous-wave distributed-feedback (DFB) quantum cascade laser (QCL) was coupled to a hollow-core optical fiber for remote mid-infrared light delivery to the test article, with high-bandwidth light detection by a direct-mounted photovoltaic detector. The method was demonstrated to measure time-resolved CO mole fraction over a range of oxidizer-to-fuel ratios and pressures (20-70 atm) in a single-element-injector RP-2-GOx rocket combustor.

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.

The ultimate limits of the relativistic rocket equation. The Planck photon rocket

NASA Astrophysics Data System (ADS)

Haug, Espen Gaarder

2017-07-01

In this paper we look at the ultimate limits of a photon propulsion rocket. The maximum velocity for a photon propulsion rocket is just below the speed of light and is a function of the reduced Compton wavelength of the heaviest subatomic particles in the rocket. We are basically combining the relativistic rocket equation with Haug's new insight on the maximum velocity for anything with rest mass. An interesting new finding is that in order to accelerate any subatomic "fundamental" particle to its maximum velocity, the particle rocket basically needs two Planck masses of initial load. This might sound illogical until one understands that subatomic particles with different masses have different maximum velocities. This can be generalized to large rockets and gives us the maximum theoretical velocity of a fully-efficient and ideal rocket. Further, no additional fuel is needed to accelerate a Planck mass particle to its maximum velocity; this also might sound absurd, but it has a very simple and logical solution that is explained in this paper.

An Italian network to improve hybrid rocket performance: Strategy and results

NASA Astrophysics Data System (ADS)

Galfetti, L.; Nasuti, F.; Pastrone, D.; Russo, A. M.

2014-03-01

The new international attention to hybrid space propulsion points out the need of a deeper understanding of physico-chemical phenomena controlling combustion process and fluid dynamics inside the motor. This research project has been carried on by a network of four Italian Universities; each of them being responsible for a specific topic. The task of Politecnico di Milano is an experimental activity concerning the study, development, manufacturing and characterization of advanced hybrid solid fuels with a high regression rate. The University of Naples is responsible for experimental activities focused on rocket motor scale characterization of the solid fuels developed and characterized at laboratory scale by Politecnico di Milano. The University of Rome has been studying the combustion chamber and nozzle of the hybrid rocket, defined in the coordinated program by advanced physical-mathematical models and numerical methods. Politecnico di Torino has been working on a multidisciplinary optimization code for optimal design of hybrid rocket motors, strongly related to the mission to be performed. The overall research project aims to increase the scientific knowledge of the combustion processes in hybrid rockets, using a strongly linked experimental-numerical approach. Methods and obtained results will be applied to implement a potential upgrade for the current generation of hybrid rocket motors. This paper presents the overall strategy, the organization, and the first experimental and numerical results of this joined effort to contribute to the development of improved hybrid propulsion systems.

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

NASA Astrophysics Data System (ADS)

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

2017-01-01

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

Solid Rocket Fuel Constitutive Theory and Polymer Cure

NASA Technical Reports Server (NTRS)

Ream, Robert

2006-01-01

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

NASP - Waveriders in a hypersonic sky.

NASA Astrophysics Data System (ADS)

Baker, David

1993-01-01

A development history is presented for the hydrogen-fueled, airbreathing (scramjet) engine-propelled National Aerospace Plane (NASP), which will be able to cruise endoatmospherically at hypersopnic speeds or rise exoatmospherically, by converting to rocket power, to LEO. Attention is given to the technology-development and configuration-validation services that the X-30 project will render the far larger NASP vehicle; the configurational and propulsion system factors in question encompass the use of 'slush' hydrogen fuel, the integration of engine inlets into the aircraft forebody and exhaust nozzles into the afterbody, and the conversion from turbojet or rocket propulsion to scramjet mode and back.

Development of high performance hybrid rocket fuels

NASA Astrophysics Data System (ADS)

Zaseck, Christopher R.

In this document I discuss paraffin fuel combustion and investigate the effects of additives on paraffin entrainment and regression. In general, hybrid rockets offer an economical and safe alternative to standard liquid and solid rockets. However, slow polymeric fuel regression and low combustion efficiency have limited the commercial use of hybrid rockets. Paraffin is a fast burning fuel that has received significant attention in the 2000's and 2010's as a replacement for standard fuels. Paraffin regresses three to four times faster than polymeric fuels due to the entrainment of a surface melt layer. However, further regression rate enhancement over the base paraffin fuel is necessary for widespread hybrid rocket adoption. I use a small scale opposed flow burner to investigate the effect of additives on the combustion of paraffin. Standard additives such as aluminum combust above the flame zone where sufficient oxidizer levels are present. As a result no heat is generated below the flame itself. In small scale opposed burner experiments the effect of limited heat feedback is apparent. Aluminum in particular does not improve the regression of paraffin in the opposed burner. The lack of heat feedback from additive combustion limits the applicability of the opposed burner. In turn, the results obtained in the opposed burner with metal additive loaded hybrid fuels do not match results from hybrid rocket experiments. In addition, nano-scale aluminum increases melt layer viscosity and greatly slows the regression of paraffin in the opposed flow burner. However, the reactive additives improve the regression rate of paraffin in the opposed burner where standard metals do not. At 5 wt.% mechanically activated titanium and carbon (Ti-C) improves the regression rate of paraffin by 47% in the opposed burner. The mechanically activated Ti C likely reacts in or near the melt layer and provides heat feedback below the flame region that results in faster opposed burner regression. In order to examine paraffin/additive combustion in a motor environment, I conducted experiments on well characterized aluminum based additives. In particular, I investigate the influence of aluminum, unpassivated aluminum, milled aluminum/polytetrafluoroethylene (PTFE), and aluminum hydride on the performance of paraffin fuels for hybrid rocket propulsion. I use an optically accessible combustor to examine the performance of the fuel mixtures in terms of characteristic velocity efficiency and regression rate. Each combustor test consumes a 12.7 cm long, 1.9 cm diameter fuel strand under 160 kg/m 2s of oxygen at up to 1.4 MPa. The experimental results indicate that the addition of 5 wt.% 30 mum or 80 nm aluminum to paraffin increases the regression rate by approximately 15% compared to neat paraffin grains. At higher aluminum concentrations and nano-scale particles sizes, the increased melt layer viscosity causes slower regression. Alane and Al/PTFE at 12.5 wt.% increase the regression of paraffin by 21% and 32% respectively. Finally, an aging study indicates that paraffin can protect air and moisture sensitive particles from oxidation. The opposed burner and aluminum/paraffin hybrid rocket experiments show that additives can alter bulk fuel properties, such as viscosity, that regulate entrainment. The general effect of melt layer properties on the entrainment and regression rate of paraffin is not well understood. Improved understanding of how solid additives affect the properties and regression of paraffin is essential to maximize performance. In this document I investigate the effect of melt layer properties on paraffin regression using inert additives. Tests are performed in the optical cylindrical combustor at ˜1 MPa under a gaseous oxygen mass flux of ˜160 kg/m2s. The experiments indicate that the regression rate is proportional to mu0.08rho 0.38kappa0.82. In addition, I explore how to predict fuel viscosity, thermal conductivity, and density prior to testing. Mechanically activated Ti-C and Al/PTFE are examined in the optical combustor. I examine the effect of the reactivity by altering the mill time for the Ti-C and Al/PTFE particles. Mechanical activation of both Ti-C and Al/PTFE improve the regression rate of paraffin more than the unmilled additives. At 12.5 wt.% Al/PTFE milled for 40 minutes regresses 12% faster than the unmilled fuel. Similarly, at 12.5 wt.% 7.5 minute milled Ti C regresses 7% faster than unmilled Ti-C. The reactive particles increase heat transfer to the fuel surface and improve regression. The composition of the combustion products are examined using a particle catcher system in conjunction with visible light and electron microscopy. The exhaust products indicate that the mechanical activation of the Al/PTFE particles cause microexplosions that reduce exhaust particle size. However, the composition of the mechanically activated Al/PTFE products do not indicate more complete combustion. In addition, the mechanically activated and unmilled Ti-C showed no difference in exhaust products.

Acoustic-Structure Interaction in Rocket Engines: Validation Testing

NASA Technical Reports Server (NTRS)

Davis, R. Benjamin; Joji, Scott S.; Parks, Russel A.; Brown, Andrew M.

2009-01-01

While analyzing a rocket engine component, it is often necessary to account for any effects that adjacent fluids (e.g., liquid fuels or oxidizers) might have on the structural dynamics of the component. To better characterize the fully coupled fluid-structure system responses, an analytical approach that models the system as a coupled expansion of rigid wall acoustic modes and in vacuo structural modes has been proposed. The present work seeks to experimentally validate this approach. To experimentally observe well-coupled system modes, the test article and fluid cavities are designed such that the uncoupled structural frequencies are comparable to the uncoupled acoustic frequencies. The test measures the natural frequencies, mode shapes, and forced response of cylindrical test articles in contact with fluid-filled cylindrical and/or annular cavities. The test article is excited with a stinger and the fluid-loaded response is acquired using a laser-doppler vibrometer. The experimentally determined fluid-loaded natural frequencies are compared directly to the results of the analytical model. Due to the geometric configuration of the test article, the analytical model is found to be valid for natural modes with circumferential wave numbers greater than four. In the case of these modes, the natural frequencies predicted by the analytical model demonstrate excellent agreement with the experimentally determined natural frequencies.

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 based on the heat flux estimates from the temperature measurements.

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

DTIC Science & Technology

2017-04-12

Clearance Number 17163 Rocket Cycles and Environments SpaceX Merlin 1D 190 klbf Russian RD-180 860 klbf Gas Generator Cycle Ox-Rich Staged Combustion...affordability & reusability • Modeling & Simulation • Key to development • Requires accurate models “CFD simulations… shorten the test-fail-fix loop” SpaceX

KSC-2011-8167

NASA Image and Video Library

2011-12-02

CAPE CANAVERAL, Fla. – A truck positions a full-size display of a space shuttle external fuel tank from the Kennedy Space Center Visitor Complex at a temporary storage area at NASA's Kennedy Space Center. The tank was part of a display of the external tank and two solid rocket boosters at the visitor complex that were used to show visitors the size of actual space shuttle components. A space shuttle rode piggyback on the tank and boosters at liftoff and during the ascent into space. The tank, which held propellants for the shuttle's three main engines, was not reused, but burned up in the atmosphere and fell into the ocean. Photo credit: NASA/Dmitri Gerondidakis

KSC-2009-5826

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – External tank 134 has arrived in the transfer aisle of the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. The Pegasus barge, carrying the fuel tank, arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

Optical Cryogenic Tank Level Sensor

NASA Technical Reports Server (NTRS)

Duffell, Amanda

2005-01-01

Cryogenic fluids play an important role in space transportation. Liquid oxygen and hydrogen are vital fuel components for liquid rocket engines. It is also difficult to accurately measure the liquid level in the cryogenic tanks containing the liquids. The current methods use thermocouple rakes, floats, or sonic meters to measure tank level. Thermocouples have problems examining the boundary between the boiling liquid and the gas inside the tanks. They are also slow to respond to temperature changes. Sonic meters need to be mounted inside the tank, but still above the liquid level. This causes problems for full tanks, or tanks that are being rotated to lie on their side.

KSC-2011-2423

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Crews in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida remove space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

KSC-2011-2425

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Crews in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida remove space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

KSC-2011-2427

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit, is removed in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

KSC-2011-2431

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit, is moved to a transporter in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

KSC-2011-2433

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit, is atop a transporter in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

KSC-2011-2430

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit, is moved to a transporter in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

KSC-2011-2426

NASA Image and Video Library

2011-03-22

CAPE CANAVERAL, Fla. - Crews in Orbiter Processing Facility-2 at NASA's Kennedy Space Center in Florida remove space shuttle Discovery's forward reaction control system (FRCS), which helped steer the shuttle in orbit. To maneuver, the FRCS used hypergolic fuel and oxidizer, which were purged from Discovery after its final spaceflight, STS-133. Next, the FRCS will be shipped to a maintenance facility at White Sands Space Harbor in New Mexico, where additional inspections will be performed and its components made safe to go on public display. The transition and retirement processing is expected to help rocket designers build next-generation spacecraft and prepare the shuttle for display. Photo credit: NASA/Jim Grossmann

Design of Force Sensor Leg for a Rocket Thrust Detector

NASA Astrophysics Data System (ADS)

Woten, Douglas; McGehee, Tripp; Wright, Anne

2005-03-01

A hybrid rocket is composed of a solid fuel and a separate liquid or gaseous oxidizer. These rockets may be throttled like liquid rockets, are safer than solid rockets, and are much less complex than liquid rockets. However, hybrid rockets produce thrust oscillations that are not practical for large scale use. A lab scale hybrid rocket at the University of Arkansas at Little Rock (UALR) Hybrid Rocket Facility is used to develop sensors to measure physical properties of hybrid rockets. Research is currently being conducted to design a six degree of freedom force sensor to measure the thrust and torque in all three spacial dimensions. The detector design uses six force sensor legs. Each leg utilizes strain gauges and a Wheatstone bridge to produce a voltage propotional to the force on the leg. The leg was designed using the CAD software ProEngineer and ProMechanica. Computer models of the strains on the single leg will be presented. A prototype leg was built and was tested in an INSTRON and results will be presented.

Astro Camp is a blast!

NASA Image and Video Library

2006-06-08

An Astro Camp counselor and her campers perform a science experiment to learn what types of `fuel' will best propel their 'rockets.' Stennis Space Center's popular series of day camps have campers design, build and test model rockets based on the principles that would be used to build different types of rockets suitable for a mission to the moon or Mars. They learn details like how far they would travel, how long it would take, what supplies they would need and how to survive in that environment.

KSC-2012-4455

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load the aft skirt for a space shuttle solid rocket booster on a truck. A twin set of space shuttle solid rocket boosters and an external fuel tank are being prepared for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

Design and testing of a caseless solid-fuel integral-rocket ramjet engine for use in small tactical missiles

NASA Astrophysics Data System (ADS)

Fruge, Keith J.

1991-09-01

An investigation was conducted to determine the feasibility of a low cost, caseless, solid fuel integral rocket ramjet (IRSFRJ) that has no ejecta. Analytical design of a ramjet powered air-to-ground missile capable of being fired from a remotely piloted vehicle or helicopter was accomplished using current JANNAF and Air Force computer codes. The results showed that an IRSFRJ powered missile can exceed the velocity and range of current systems by more than a two to one ratio, without an increase in missile length and weight. A caseless IRSFRJ with a nonejecting port cover was designed and tested. The experimental results of the static tests showed that a low cost, caseless IRSFRJ with a nonejectable port cover is a viable design. Rocket ramjet transition was demonstrated and ramjet ignition was found to be insensitive to the booster tail off to air injection timing sequence.

Current Development of Nuclear Thermal Propulsion technologies at the Center for Space Nuclear Research

DOE Office of Scientific and Technical Information (OSTI.GOV)

Robert C. O'Brien; Steven K. Cook; Nathan D. Jerred

Nuclear power and propulsion has been considered for space applications since the 1950s. Between 1955 and 1972 the US built and tested over twenty nuclear reactors / rocket engines in the Rover/NERVA programs1. The Aerojet Corporation was the prime contractor for the NERVA program. Modern changes in environmental laws present challenges for the redevelopment of the nuclear rocket. Recent advances in fuel fabrication and testing options indicate that a nuclear rocket with a fuel composition that is significantly different from those of the NERVA project can be engineered; this may be needed to ensure public support and compliance with safetymore » requirements. The Center for Space Nuclear Research (CSNR) is pursuing a number of technologies, modeling and testing processes to further the development of safe, practical and affordable nuclear thermal propulsion systems.« less

Space Shuttle Projects

NASA Image and Video Library

2004-04-15

The Apollo program demonstrated that men could travel into space, perform useful tasks there, and return safely to Earth. But space had to be more accessible. This led to the development of the Space Shuttle. The Shuttle's major components are the orbiter spacecraft; the three main engines, with a combined thrust of more than 1.2 million pounds; the huge external tank (ET) that feeds the liquid hydrogen fuel and liquid oxygen oxidizer to the three main engines; and the two solid rocket boosters (SRBs), with their combined thrust of some 5.8 million pounds, that provide most of the power for the first two minutes of flight. Crucially involved with the Space Shuttle program virtually from its inception, the Marshall Space Flight Center (MSFC) played a leading role in the design, development, testing, and fabrication of many major Shuttle propulsion components.

Space Shuttle Drawing

NASA Technical Reports Server (NTRS)

2004-01-01

The Apollo program demonstrated that men could travel into space, perform useful tasks there, and return safely to Earth. But space had to be more accessible. This led to the development of the Space Shuttle. The Shuttle's major components are the orbiter spacecraft; the three main engines, with a combined thrust of more than 1.2 million pounds; the huge external tank (ET) that feeds the liquid hydrogen fuel and liquid oxygen oxidizer to the three main engines; and the two solid rocket boosters (SRBs), with their combined thrust of some 5.8 million pounds, that provide most of the power for the first two minutes of flight. Crucially involved with the Space Shuttle program virtually from its inception, the Marshall Space Flight Center (MSFC) played a leading role in the design, development, testing, and fabrication of many major Shuttle propulsion components.

KSC-2009-1577

NASA Image and Video Library

2009-02-12

CAPE CANAVERAL, Fla. – An aerial view shows the entire Kennedy Space Center Visitor Complex in Florida. In the foreground at left are a full-size orbiter replica, the "Explorer," which allows guests to picture more clearly what it's like to live and work in space, as well as full-size, genuine solid rocket boosters and an external fuel tank (foreground, left). On the right at center is the rocket garden, which features eight authentic rockets from the past, including a Mercury-Atlas rocket similar to the one used to launch John Glenn into space in 1962 and climb-in Mercury, Gemini and Apollo capsule rep¬licas. Photo credit: NASA/Kim Shiflett

Karl Poggensee - A widely unknown German rocket pioneer - The early years 1930-1934 - A chronology

NASA Astrophysics Data System (ADS)

Rohrwild, Karlheinz

2017-09-01

The rediscovered estate of Karl Poggensee allows to reproduce chronologically his rocket tests of the period 1930-1934 almost completely for the first time. Thrilled by the movie ;The Woman in the Moon; for the idea of space travel, he started as a student of Hinderburg-Polytechnikum (IAO), Oldenburg, to build his first solid-fuel rocket, producing his own propellant charges. Being a coming electrical engineer his main goal was not set up new record heights, but to provide his rockets with automatic measuring instruments, camera and parachute release systems. The optimization of this sequence was his main focus.

Deposit formation and heat transfer in hydrocarbon rocket fuels

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

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

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

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

Space Shuttle Transportation (Roll-Out) Loads Diagnostics

NASA Technical Reports Server (NTRS)

Elliott, Kenny B.; Buehrle, Ralph D.; James, George H.; Richart, Jene A.

2005-01-01

The Space Transportation System (STS) consists of three primary components; an Orbiter Vehicle, an External Fuel Tank, and two Solid Rocket Boosters. The Orbiter Vehicle and Solid Rocket Boosters are reusable components, and as such, they are susceptible to durability issues. Recently, the fatigue load spectra for these components have been updated to include load histories acquired during the rollout phase of the STS processing for flight. Using traditional program life assessment techniques, the incorporation of these "rollout" loads produced unacceptable life estimates for certain Orbiter structural members. As a result, the Space Shuttle System Engineering and Integration Office has initiated a program to re-assess the method used for developing the "rollout" loads and performing the life assessments. In the fall of 2003 a set of tests were preformed to provide information to either validate existing load spectra estimation techniques or generate new load spectra estimation methods. Acceleration and strain data were collected from two rollouts of a partial-stack configuration of the Space Shuttle. The partial stack configuration consists of two Solid Rocket Boosters tied together at the upper External Tank attachment locations mounted on the Mobile Launch Platform carried by a Crawler Transporter (CT). In the current analysis, the data collected from this test is examined for consistency in speed, surface condition effects, and the characterization of the forcing function. It is observed that the speed of the CT is relatively stable. The dynamic response acceleration of the partial-stack is slightly sensitive to the surface condition of the road used for transport, and the dynamic response acceleration of the partial-stack generally increases as the transport speed increases. However, the speed sensitivity is dependent on the measurement location. Finally, the character of the forcing function is narrow-banded with the primary drivers being harmonics of two CT speed dependent excitations. One source is an excitation due to the CT treads striking the road surface, and the second is unknown.

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.

Material Science of Carbon

DTIC Science & Technology

2004-09-01

required for a specific application. The list of applications is very extensive and includes: aircraft brakes, electrodes, high temperature molds, rocket...and includes: aircraft brakes, electrodes, high temperature molds, rocket nozzles and exit cones, tires, ink, nuclear reactors and fuel particles...produced. For example carbons can be hard (chars) or soft (blacks), strong (PAN fibers) or weak ( aerogel ), stiff (pitch fibers) or flexible

Program For Optimization Of Nuclear Rocket Engines

NASA Technical Reports Server (NTRS)

Plebuch, R. K.; Mcdougall, J. K.; Ridolphi, F.; Walton, James T.

1994-01-01

NOP is versatile digital-computer program devoloped for parametric analysis of beryllium-reflected, graphite-moderated nuclear rocket engines. Facilitates analysis of performance of engine with respect to such considerations as specific impulse, engine power, type of engine cycle, and engine-design constraints arising from complications of fuel loading and internal gradients of temperature. Predicts minimum weight for specified performance.

Around Marshall

NASA Image and Video Library

2002-10-01

This is a ground level view of Test Stand 300 at the east test area of the Marshall Space Flight Center. Test Stand 300 was constructed in 1964 as a gas generator and heat exchanger test facility to support the Saturn/Apollo Program. Deep-space simulation was provided by a 1960 modification that added a 20-ft thermal vacuum chamber and a 1981 modification that added a 12-ft vacuum chamber. The facility was again modified in 1989 when 3-ft and 15-ft diameter chambers were added to support Space Station and technology programs. This multiposition test stand is used to test a wide range of rocket engine components, systems, and subsystems. It has the capability to simulate launch thermal and pressure profiles. Test Stand 300 was designed for testing solid rocket booster (SRB) insulation panels and components, super-insulated tanks, external tank (ET) insulation panels and components, Space Shuttle components, solid rocket motor materials, and advanced solid rocket motor materials.

Prediction of high frequency combustion instability in liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

Kim, Y. M.; Chen, C. P.; Ziebarth, J. P.; Chen, Y. S.

1992-01-01

The present use of a numerical model developed for the prediction of high-frequency combustion stabilities in liquid propellant rocket engines focuses on (1) the overall behavior of nonlinear combustion instabilities (2) the effects of acoustic oscillations on the fuel-droplet vaporization and combustion process in stable and unstable engine operating conditions, oscillating flowfields, and liquid-fuel trajectories during combustion instability, and (3) the effects of such design parameters as inlet boundary conditions, initial spray conditions, and baffle length. The numerical model has yielded predictions of the tangential-mode combustion instability; baffle length and droplet size variations are noted to have significant effects on engine stability.

Ground test facility for SEI nuclear rocket engines

NASA Astrophysics Data System (ADS)

Harmon, Charles D.; Ottinger, Cathy A.; Sanchez, Lawrence C.; Shipers, Larry R.

1992-07-01

Nuclear (fission) thermal propulsion has been identified as a critical technology for a manned mission to Mars by the year 2019. Facilities are required that will support ground tests to qualify the nuclear rocket engine design, which must support a realistic thermal and neutronic environment in which the fuel elements will operate at a fraction of the power for a flight weight reactor/engine. This paper describes the design of a fuel element ground test facility, with a strong emphasis on safety and economy. The details of major structures and support systems of the facility are discussed, and a design diagram of the test facility structures is presented.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

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

2010-05-01

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

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.

The design and evolution of the beta two-stage-to-orbit horizontal takeoff and landing launch system

NASA Technical Reports Server (NTRS)

Burkardt, Leo A.; Norris, Rick B.

1992-01-01

The Beta launch system was originally conceived in 1986 as a horizontal takeoff and landing, fully reusable, two-stage-to-orbit, manned launch vehicle to replace the Shuttle. It was to be capable of delivering a 50,000 lb. payload to low polar orbit. The booster propulsion system consisted of JP fueled turbojets and LH fueled ramjets mounted in pods in an over/under arrangement, and a single LOX/LH fueled SSME rocket. The second stage orbiter, which staged at Mach 8, was powered by an SSME rocket. A major goal was to develop a vehicle design consistent with near term technology. The vehicle design was completed with a GLOW of approximately 2,000,000 lbs. All design goals were met. Since then, interest has shifted to the 10,000 lbs. to low polar orbit payload class. The original Beta was down-sized to meet this payload class. The GLOW of the down-sized vehicle was approximately 1,000,000 lbs. The booster was converted to exclusively air-breathing operation. Because the booster depends on conventional air-breathing propulsion only, the staging Mach number was reduced to 5.5. The orbiter remains an SSME rocket-powered stage.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zittel, P.F.

1994-09-10

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

Launch Vehicles

NASA Image and Video Library

2007-09-09

Under the goals of the Vision for Space Exploration, Ares I is a chief component of the cost-effective space transportation infrastructure being developed by NASA's Constellation Program. This transportation system will safely and reliably carry human explorers back to the moon, and then onward to Mars and other destinations in the solar system. The Ares I effort includes multiple project element teams at NASA centers and contract organizations around the nation, and is managed by the Exploration Launch Projects Office at NASA's Marshall Space Flight Center (MFSC). ATK Launch Systems near Brigham City, Utah, is the prime contractor for the first stage booster. ATK's subcontractor, United Space Alliance of Houston, is designing, developing and testing the parachutes at its facilities at NASA's Kennedy Space Center in Florida. NASA's Johnson Space Center in Houston hosts the Constellation Program and Orion Crew Capsule Project Office and provides test instrumentation and support personnel. Together, these teams are developing vehicle hardware, evolving proven technologies, and testing components and systems. Their work builds on powerful, reliable space shuttle propulsion elements and nearly a half-century of NASA space flight experience and technological advances. Ares I is an inline, two-stage rocket configuration topped by the Crew Exploration Vehicle, its service module, and a launch abort system. The launch vehicle's first stage is a single, five-segment reusable solid rocket booster derived from the Space Shuttle Program's reusable solid rocket motor that burns a specially formulated and shaped solid propellant called polybutadiene acrylonitrile (PBAN). The second or upper stage will be propelled by a J-2X main engine fueled with liquid oxygen and liquid hydrogen. This HD video image depicts a test firing of a 40k subscale J2X injector at MSFC's test stand 115. (Highest resolution available)

Scout

NASA Image and Video Library

1960-09-22

Photographed on 09/22/1960. -- An examination of the Aerojet-General "Aerobee 150A" propulsion system in February 1960. James Hansen described this as follows: "As for the technical definition of the rocket...the Langley engineers tried to keep developmental costs and time to a minimum by selecting components from off-the-shelf hardware. the majority of Scout's components were to come from an inventory of solid-fuel rockets produced for the military, although everyone involved understood that some improved motors would also have to be developed under contract. By early 1959, after intensive technical analysis and reviews, Langley settled on a design and finalized the selection of the major contractors. The rocket's 40-inch-diameter first stage was to be a new "Algol" motor, a combination of the Jupiter Senior and the navy Polaris produced by the Aerojet General Corporation, Sacramento, California. The 31-inch-diameter second stage, "Castor," was derived from the army's Sergeant and was to be manufactured by the Redstone Division of the Thiokol company in Huntsville, Alabama. the motor for the 30-inch-diameter third stage, "Antares," evolved under NASA contract from the ABL X248 design into a new version called the X254 (and subsequently into the X259); it was built under contract to NASA by ABL, a U.S. Navy Bureau of Ordnance facility operated by the Hercules Powder Company, Cumberland, Maryland. the final upper-stage propulsion unit, "Altair," which was 25.7 inches in diameter (34 inches at the heat shield), amounted to an improved edition of the X248 that was also manufactured by ABL." -- Published in James R. Hansen, Spaceflight Revolution: NASA Langley Research Center From Sputnik to Apollo, NASA SP-4308, pp.200-201.

Design of a 500 lbf liquid oxygen and liquid methane rocket engine for suborbital flight

NASA Astrophysics Data System (ADS)

Trillo, Jesus Eduardo

Liquid methane (LCH4)is the most promising rocket fuel for our journey to Mars and other space entities. Compared to liquid hydrogen, the most common cryogenic fuel used today, methane is denser and can be stored at a more manageable temperature; leading to more affordable tanks and a lighter system. The most important advantage is it can be produced from local sources using in-situ resource utilization (ISRU) technology. This will allow the production of the fuel needed to come back to earth on the surface of Mars, or the space entity being explored, making the overall mission more cost effective by enabling larger usable mass. The major disadvantage methane has over hydrogen is it provides a lower specific impulse, or lower rocket performance. The UTEP Center for Space Exploration and Technology Research (cSETR) in partnership with the National Aeronautics and Space Administration (NASA) has been the leading research center for the advancement of Liquid Oxygen (LOX) and Liquid Methane (LCH4) propulsion technologies. Through this partnership, the CROME engine, a throattable 500 lbf LOX/LCH4 rocket engine, was designed and developed. The engine will serve as the main propulsion system for Daedalus, a suborbital demonstration vehicle being developed by the cSETR. The purpose of Daedalus mission and the engine is to fire in space under microgravity conditions to demonstrate its restartability. This thesis details the design process, decisions, and characteristics of the engine to serve as a complete design guide.

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

NASA Technical Reports Server (NTRS)

Ladanyi, Dezso J

1952-01-01

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

Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation

NASA Technical Reports Server (NTRS)

Sivo, Joseph N.; Peters, Daniel J.

1959-01-01

A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.

48 CFR 9.508 - Examples.

Code of Federal Regulations, 2012 CFR

2012-10-01

... contract to define the detailed performance characteristics an agency will require for purchasing rocket fuels. Company A has not developed the particular fuels. When the definition contract is awarded, it is clear to both parties that the agency will use the performance characteristics arrived at to choose...

48 CFR 9.508 - Examples.

Code of Federal Regulations, 2013 CFR

2013-10-01

... contract to define the detailed performance characteristics an agency will require for purchasing rocket fuels. Company A has not developed the particular fuels. When the definition contract is awarded, it is clear to both parties that the agency will use the performance characteristics arrived at to choose...

48 CFR 9.508 - Examples.

Code of Federal Regulations, 2014 CFR

2014-10-01

... contract to define the detailed performance characteristics an agency will require for purchasing rocket fuels. Company A has not developed the particular fuels. When the definition contract is awarded, it is clear to both parties that the agency will use the performance characteristics arrived at to choose...

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

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

Report on Hybrid Rocket Cold Flow Experiments

NASA Technical Reports Server (NTRS)

Haapanen, Siina

2004-01-01

The discovery of paraffin based fuels has lead to renewed interest in hybrid rocket research. Experiments have shown that they burn 3-5 times faster than conventional hybrid fuels. High thrust level that would have required a multi-port design in the past can now be achieved with a single-port motor. While tests performed in Stanford and NASA Ames have demonstrated the paraffin hybrids to be a promising technology, one of the major challenges has been the relatively low efficiency. The c* efficiency has ranged between 80% and 90% in experiments conducted at the Ames Hybrid Combustion Facility (HCF). The test motor in these experiments had a 45 inch long fuel grain with the initial port diameter ranging between 3 and 5_inches. The c* efficiency is defined as the ratio of measured and theoretical characteristic velocities and is related to how completely the fuel and oxidizer are converted to combustion products. A low efficiency means that the reactants burn incompletely, and the reaction does not release the maximum possible amount of energy.

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

NASA Technical Reports Server (NTRS)

Martin, J. A.

1978-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

An Ignition Torch Based on Photoignition of Carbon Nanotubes at Elevated Pressure (Briefing Charts)

DTIC Science & Technology

2016-01-04

Ignition Capsule A 10 mg low pressure ignition torch as it ignites a fuel spray We use PITCH to ignite subscale test rockets at 130 K and ~35 atm (~500...distribution is unlimited High Pressure PITCH Applied to a H2/O2 Subscale Rocket Injector Top: a high-pressure chamber for test of subscale rocket injector...to a high-pressure test combustion chamber via a 20 cm extension tube (OD=6 mm) Click >>> 9 DISTRIBUTION STATEMENT A. Approved for public release

Hybrid boosters for future launch vehicles

NASA Astrophysics Data System (ADS)

Dargies, E.; Lo, R. E.

1987-10-01

Hybrid rocket propulsion systems furnish the advantages of much higher safety levels, due both to shut-down capability in case of ignition failure to one unit and the potential choice of nontoxic propellant combinations, such as LOX/polyethylene; they nevertheless yield performance levels comparable or superior to those of solid rocket boosters. Attention is presently given to the results of DFVLR analytical model studies of hybrid propulsion systems, with attention to solid fuel grain geometrical design and propellant grain surface ablation rate. The safety of hybrid rockets recommends them for use by manned spacecraft.

Animation: What makes up the Space Launch System’s massive core stage

NASA Image and Video Library

2017-04-24

NASA’s new rocket, the Space Launch System, will be the most powerful rocket ever built for deep-space missions. The 212-foot core stage is the largest rocket stage ever built and will fuel four RS-25 engines that will help launch SLS. This animation depicts the parts that make up the core stage and how these parts will be joined to form the entire stage. The five major parts include: the engine section, the hydrogen tank, the intertank, the liquid oxygen tank and the forward skirt.

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

NASA Image and Video Library

2016-08-18

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

Solid rocket motors for the Space Shuttle booster.

NASA Technical Reports Server (NTRS)

Odom, J. B.

1972-01-01

The evolution of the space shuttle booster system is reviewed from its initial concepts based on liquid-propellant reusable boosters to the final selection of recoverable, solid-fuel rocket motors. The rationale associated with each of the several major decisions in the evolution process is discussed. It is shown that the external tank orbiter configuration emerging from the latest studies takes maximum advantage of the solid rocket motor development experience and promises to be the optimum configuration for fulfilling the paramount shuttle program requirements of minimum total development risk within acceptable costs.

Launch Vehicles Based on Advanced Hybrid Rocket Motors: An Enabling Technology for the Commercial Small and Micro Satellite Planetary Science

NASA Astrophysics Data System (ADS)

Karabeyoglu, Arif; Tuncer, Onur; Inalhan, Gokhan

2016-07-01

Mankind is relient on chemical propulsion systems for space access. Nevertheless, this has been a stagnant area in terms of technological development and the technology base has not changed much almost for the past forty years. This poses a vicious circle for launch applications such that high launch costs constrain the demand and low launch freqencies drive costs higher. This also has been a key limiting factor for small and micro satellites that are geared towards planetary science. Rather this be because of the launch frequencies or the costs, the access of small and micro satellites to orbit has been limited. With today's technology it is not possible to escape this circle. However the emergence of cost effective and high performance propulsion systems such as advanced hybrid rockets can decrease launch costs by almost an order or magnitude. This paper briefly introduces the timeline and research challenges that were overcome during the development of advanced hybrid LOX/paraffin based rockets. Experimental studies demonstrated effectiveness of these advanced hybrid rockets which incorporate fast burning parafin based fuels, advanced yet simple internal balistic design and carbon composite winding/fuel casting technology that enables the rocket motor to be built from inside out. A feasibility scenario is studied using these rocket motors as building blocks for a modular launch vehicle capable of delivering micro satellites into low earth orbit. In addition, the building block rocket motor can be used further solar system missions providing the ability to do standalone small and micro satellite missions to planets within the solar system. This enabling technology therefore offers a viable alternative in order to escape the viscous that has plagued the space launch industry and that has limited the small and micro satellite delivery for planetary science.

n/a

NASA Image and Video Library

1963-03-28

The Saturn I (SA-4) flight lifted off from Kennedy Space Center launch Complex 34, March 28, 1963. The fourth launch of Saturn launch vehicles developed at the Marshall Space Flight Center (MSFC), under the direction of Dr. Wernher von Braun, incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle. Like SA-3, the SA-4 flight’s upper stage ejected 113,560 liters (30,000 gallons) of ballast water in the upper atmosphere for "Project Highwater" physics experiment. Release of this vast quantity of water in a near-space environment marked the second purely scientific large-scale experiment. The SA-4 was the last Block I rocket launch.

n/a

NASA Image and Video Library

1963-03-28

The Saturn I (SA-4) flight lifted off from Kennedy Space Center launch Complex 34, March 28, 1963. The fourth launch of Saturn launch vehicles, developed at the Marshall Space Flight Center (MSFC) under the direction of Dr. Wernher von Braun, incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle. Like SA-3, the SA-4 flight’s upper stage ejected 113,560 liters (30,000 gallons) of ballast water in the upper atmosphere for "Project Highwater" physics experiment. Release of this vast quantity of water in a near-space environment marked the second purely scientific large-scale experiment. The SA-4 was the last Block I rocket launch.

U.S./CIS eye joint nuclear rocket venture

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

[Integral evaluation of immune homeostasis in rockets liquidators and role of this evaluation for prophylaxis].

PubMed

2010-01-01

Long-standing clinical and immunologic monitoring and integral evaluation of immune homeostasis (through generalized parameter) in personnel of Center for liquid-fuel rockets liquidation demonstrated diagnostically reliable immunity parameters that enable to forecast changes in the workers' health state. The authors defined boundary values of the generalized parameter to form risk groups for specific entities formation.

The 17th JANNAF Combustion Meeting, Volume 1

NASA Technical Reports Server (NTRS)

Eggleston, D. S. (Editor)

1980-01-01

The combustion of solid rocket propellants and combustion in ramjets is addressed. Subjects discussed include metal burning, steady-state combustion of composite propellants, velocity coupling and nonlinear instability, vortex shedding and flow effects on combustion instability, combustion instability in solid rocket motors, combustion diagnostics, subsonic and supersonic ramjet combustion, characterization of ramburner flowfields, and injection and combustion of ramjet fuels.

Modeling Primary Atomization Processes

DTIC Science & Technology

1999-02-01

consumable , catalytic igniter has shown to provide reliable, reproducible ignition in hydrogen peroxide/polyethylene hybrid engines. Currently, a...verified in a hybrid rocket using hydrogen peroxide as oxidizer and polyethylene as fuel. The engine made use of a unique Consumable Catalytic Bed (CCB...interest to the liquid and hybrid rocket engine community. TECHNOLOGY TRANSFER Performer Customer Result Application 1 S. D. Heister Purdue University

Multiple-wavelength transmission measurements in rocket motor plumes

NASA Astrophysics Data System (ADS)

Kim, Hong-On

1991-09-01

Multiple-wavelength light transmission measurements were used to measure the mean particle size (d(sub 32)), index of refraction (m), and standard deviation of the small particles in the edge of the plume of a small solid propellant rocket motor. The results have shown that the multiple-wavelength light transmission measurement technique can be used to obtain these variables. The technique was shown to be more sensitive to changes in d(sub 32) and standard deviation (sigma) than to m. A GAP/AP/4.7 percent aluminum propellant burned at 25 atm produced particles with d32 = 0.150 +/- 0.006 microns, standard deviation = 1.50 +/- 0.04 and m = 1.63 +/- 0.13. The good correlation of the data indicated that only submicron particles were present in the edge of the plume. In today's budget conscious industry, the solid propellant rocket motor is an ideal propulsion system due to its low cost and simplicity. The major obstacle for solid rocket motors, however, is their limited specific impulse compared to airbreathing motors. One way to help overcome this limitation is to utilize metal fuel additives. Solid propellant rocket motors can achieve high specific impulse with metal fuel additives such as aluminum. Aluminum propellants also increase propellant densities and suppress transverse modes of combustion oscillations by damping the oscillations with the aluminum agglomerates in the combustion chamber.

Combustion Processes in Hybrid Rocket Engines

NASA Technical Reports Server (NTRS)

Venkateswaran,S.; Merkle, C. L.

1996-01-01

In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. Hybrid propulsion systems use a solid fuel such as hydroxyl-terminated polybutadiene (HTPB) along with a gaseous/liquid oxidizer. The performance of hybrid combustors depends on the convective and radiative heat fluxes to the fuel surface, the rate of pyrolysis in the solid phase, and the turbulent combustion processes in the gaseous phases. These processes in combination specify the regression rates of the fuel surface and thereby the utilization efficiency of the fuel. In this paper, we employ computational fluid dynamics (CFD) techniques in order to gain a quantitative understanding of the physical trends in hybrid rocket combustors. The computational modeling is tailored to ongoing experiments at Penn State that employ a two dimensional slab burner configuration. The coordinated computational/experimental effort enables model validation while providing an understanding of the experimental observations. Computations to date have included the full length geometry with and with the aft nozzle section as well as shorter length domains for extensive parametric characterization. HTPB is sed as the fuel with 1,3 butadiene being taken as the gaseous product of the pyrolysis. Pure gaseous oxygen is taken as the oxidizer. The fuel regression rate is specified using an Arrhenius rate reaction, which the fuel surface temperature is given by an energy balance involving gas-phase convection and radiation as well as thermal conduction in the solid-phase. For the gas-phase combustion, a two step global reaction is used. The standard kappa - epsilon model is used for turbulence closure. Radiation is presently treated using a simple diffusion approximation which is valid for large optical path lengths, representative of radiation from soot particles. Computational results are obtained to determine the trends in the fuel burning or regression rates as a function of the head-end oxidizer mass flux, G=rho(e)U(e), and the chamber pressure. Furthermore, computation of the full slab burner configuration has also been obtained for various stages of the burn. Comparisons with available experimental data from small scale tests conducted by General Dynamics-Thiokol-Rocketdyne suggest reasonable agreement in the predicted regression rates. Future work will include: (1) a model for soot generation in the flame for more quantitative radiative transfer modelling, (2) a parametric study of combustion efficiency, and (3) transient calculations to help determine the possible mechanisms responsible for combustion instability in hybrid rocket motors.

Critical Review of the Navy Space Cadre

DTIC Science & Technology

2014-06-01

SPACE PROGRAM Following World War II, Army and Navy researchers divided captured German V-2 rocket components to rebuild V-2 rockets and eventually...develop the first American rockets .19 The Navy began launching space probes on V-2 rockets in 1946, including from the deck of USS Midway (CVB-41) in...1947.20 But the dwindling supply of V-2 rockets motivated the Navy to develop its own rockets . The Aerobee and Viking, would form a solid foundation

An Object Model for a Rocket Engine Numerical Simulator

NASA Technical Reports Server (NTRS)

Mitra, D.; Bhalla, P. N.; Pratap, V.; Reddy, P.

1998-01-01

Rocket Engine Numerical Simulator (RENS) is a packet of software which numerically simulates the behavior of a rocket engine. Different parameters of the components of an engine is the input to these programs. Depending on these given parameters the programs output the behaviors of those components. These behavioral values are then used to guide the design of or to diagnose a model of a rocket engine "built" by a composition of these programs simulating different components of the engine system. In order to use this software package effectively one needs to have a flexible model of a rocket engine. These programs simulating different components then should be plugged into this modular representation. Our project is to develop an object based model of such an engine system. We are following an iterative and incremental approach in developing the model, as is the standard practice in the area of object oriented design and analysis of softwares. This process involves three stages: object modeling to represent the components and sub-components of a rocket engine, dynamic modeling to capture the temporal and behavioral aspects of the system, and functional modeling to represent the transformational aspects. This article reports on the first phase of our activity under a grant (RENS) from the NASA Lewis Research center. We have utilized Rambaugh's object modeling technique and the tool UML for this purpose. The classes of a rocket engine propulsion system are developed and some of them are presented in this report. The next step, developing a dynamic model for RENS, is also touched upon here. In this paper we will also discuss the advantages of using object-based modeling for developing this type of an integrated simulator over other tools like an expert systems shell or a procedural language, e.g., FORTRAN. Attempts have been made in the past to use such techniques.

KSC-2009-5823

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – External tank 134 is towed toward the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. The Pegasus barge, carrying the fuel tank, arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. Next, the tank will be transported into the VAB where it will be stored until needed. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5822

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – External tank 134 emerges from the Pegasus barge docked in the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. The fuel tank next will be transported into the VAB where it will be stored until needed. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

Enrichment Zoning Options for the Small Nuclear Rocket Engine (SNRE)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bruce G. Schnitzler; Stanley K. Borowski

2010-07-01

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. In NASA’s recent Mars Design Reference Architecture (DRA) 5.0 study (NASA-SP-2009-566, July 2009), nuclear thermal propulsion (NTP) was again selected over chemical propulsion as the preferred in-space transportation system option because of its high thrust and high specific impulse (-900 s) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit. An extensive nuclear thermal rocket technology development effortmore » was conducted from 1955-1973 under the Rover/NERVA Program. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art design incorporating lessons learned from the very successful technology development program. Past activities at the NASA Glenn Research Center have included development of highly detailed MCNP Monte Carlo transport models of the SNRE and other small engine designs. Preliminary core configurations typically employ fuel elements with fixed fuel composition and fissile material enrichment. Uniform fuel loadings result in undesirable radial power and temperature profiles in the engines. Engine performance can be improved by some combination of propellant flow control at the fuel element level and by varying the fuel composition. Enrichment zoning at the fuel element level with lower enrichments in the higher power elements at the core center and on the core periphery is particularly effective. Power flattening by enrichment zoning typically results in more uniform propellant exit temperatures and improved engine performance. For the SNRE, element enrichment zoning provided very flat radial power profiles with 551 of the 564 fuel elements within 1% of the average element power. Results for this and alternate enrichment zoning options for the SNRE are compared.« less

The Manufacture of W-UO2 Fuel Elements for NTP Using the Hot Isostatic Pressing Consolidation Process

NASA Technical Reports Server (NTRS)

Broadway, Jeramie; Hickman, Robert; Mireles, Omar

2012-01-01

NTP is attractive for space exploration because: (1) Higher Isp than traditional chemical rockets (2)Shorter trip times (3) Reduced propellant mass (4) Increased payload. Lack of qualified fuel material is a key risk (cost, schedule, and performance). Development of stable fuel form is a critical path, long lead activity. Goals of this project are: Mature CERMET and Graphite based fuel materials and Develop and demonstrate critical technologies and capabilities.

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket, 1958-2002

NASA Technical Reports Server (NTRS)

Dawson, Virginia P.; Bowles, Mark D.

2004-01-01

During its maiden voyage in May 1962, a Centaur upper stage rocket, mated to an Atlas booster, exploded 54 seconds after launch, engulfing the rocket in a huge fireball. Investigation revealed that Centaur's light, stainless-steel tank had split open, spilling its liquid-hydrogen fuel down its sides, where the flame of the rocket exhaust immediately ignited it. Coming less than a year after President Kennedy had made landing human beings on the Moon a national priority, the loss of Centaur was regarded as a serious setback for the National Aeronautics and Space Administration (NASA). During the failure investigation, Homer Newell, Director of Space Sciences, ruefully declared: "Taming liquid hydrogen to the point where expensive operational space missions can be committed to it has turned out to be more difficult than anyone supposed at the outset." After this failure, Centaur critics, led by Wernher von Braun, mounted a campaign to cancel the program. In addition to the unknowns associated with liquid hydrogen, he objected to the unusual design of Centaur. Like the Atlas rocket, Centaur depended on pressure to keep its paper-thin, stainless-steel shell from collapsing. It was literally inflated with its propellants like a football or balloon and needed no internal structure to give it added strength and stability. The so-called "pressure-stabilized structure" of Centaur, coupled with the light weight of its high- energy cryogenic propellants, made Centaur lighter and more powerful than upper stages that used conventional fuel. But, the critics argued, it would never become the reliable rocket that the United States needed.

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

NASA Technical Reports Server (NTRS)

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

1950-01-01

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

Computer program for prediction of fuel consumption statistical data for an upper stage three-axes stabilized on-off control system

NASA Technical Reports Server (NTRS)

1982-01-01

A FORTRAN coded computer program and method to predict the reaction control fuel consumption statistics for a three axis stabilized rocket vehicle upper stage is described. A Monte Carlo approach is used which is more efficient by using closed form estimates of impulses. The effects of rocket motor thrust misalignment, static unbalance, aerodynamic disturbances, and deviations in trajectory, mass properties and control system characteristics are included. This routine can be applied to many types of on-off reaction controlled vehicles. The pseudorandom number generation and statistical analyses subroutines including the output histograms can be used for other Monte Carlo analyses problems.

Rho-Isp Revisited and Basic Stage Mass Estimating for Launch Vehicle Conceptual Sizing Studies

NASA Technical Reports Server (NTRS)

Kibbey, Timothy P.

2015-01-01

The ideal rocket equation is manipulated to demonstrate the essential link between propellant density and specific impulse as the two primary stage performance drivers for a launch vehicle. This is illustrated by examining volume-limited stages such as first stages and boosters. This proves to be a good approximation for first-order or Phase A vehicle design studies for solid rocket motors and for liquid stages, except when comparing to hydrogen-fueled stages. A next-order mass model is developed that is able to model the mass differences between hydrogen-fueled and other stages. Propellants considered range in density from liquid methane to inhibited red fuming nitric acid. Calculated comparisons are shown for solid rocket boosters, liquid first stages, liquid upper stages, and a balloon-deployed single-stage-to-orbit concept. The derived relationships are ripe for inclusion in a multi-stage design space exploration and optimization algorithm, as well as for single-parameter comparisons such as those shown herein.

RTE: A computer code for Rocket Thermal Evaluation

NASA Technical Reports Server (NTRS)

Naraghi, Mohammad H. N.

1995-01-01

The numerical model for a rocket thermal analysis code (RTE) is discussed. RTE is a comprehensive thermal analysis code for thermal analysis of regeneratively cooled rocket engines. The input to the code consists of the composition of fuel/oxidant mixture and flow rates, chamber pressure, coolant temperature and pressure. dimensions of the engine, materials and the number of nodes in different parts of the engine. The code allows for temperature variation in axial, radial and circumferential directions. By implementing an iterative scheme, it provides nodal temperature distribution, rates of heat transfer, hot gas and coolant thermal and transport properties. The fuel/oxidant mixture ratio can be varied along the thrust chamber. This feature allows the user to incorporate a non-equilibrium model or an energy release model for the hot-gas-side. The user has the option of bypassing the hot-gas-side calculations and directly inputting the gas-side fluxes. This feature is used to link RTE to a boundary layer module for the hot-gas-side heat flux calculations.

Towards Flange-to-Flange Turbopump Simulations for Liquid Rocket Engines

NASA Technical Reports Server (NTRS)

Kiris, Cetin; Williams, Robert

2000-01-01

The primary objective of this research is to support the design of liquid rocket systems for the Advanced Space Transportation System. Since the space launch systems in the near future are likely to rely on liquid rocket engines, increasing the efficiency and reliability of the engine components is an important task. One of the major problems in the liquid rocket engine is to understand fluid dynamics of fuel and oxidizer flows from the fuel tank to plume. Understanding the flow through the entire turbopump geometry through numerical simulation will be of significant value toward design. This will help to improve safety of future space missions. One of the milestones of this effort is to develop, apply and demonstrate the capability and accuracy of 3D CFD methods as efficient design analysis tools on high performance computer platforms. The development of the MPI and MLP versions of the INS3D code is currently underway. The serial version of INS3D code is a multidimensional incompressible Navier-Stokes solver based on overset grid technology. INS3D-MPI is based on the explicit massage-passing interface across processors and is primarily suited for distributed memory systems. INS3D-MLP is based on multi-level parallel method and is suitable for distributed-shared memory systems. For the entire turbopump simulations, moving boundary capability and an efficient time-accurate integration methods are build in the flow solver. To handle the geometric complexity and moving boundary problems, overset grid scheme is incorporated with the solver that new connectivity data will be obtained at each time step. The Chimera overlapped grid scheme allows subdomains move relative to each other, and provides a great flexibility when the boundary movement creates large displacements. The performance of the two time integration schemes for time-accurate computations is investigated. For an unsteady flow which requires small physical time step, the pressure projection method was found to be computationally efficient since it does not require any subiterations procedure. It was observed that the artificial compressibility method requires a fast convergence scheme at each physical time step in order to satisfy incompressibility condition. This was obtained by using a GMRES-ILU(0) solver in our computations. When a line-relaxation scheme was used, the time accuracy was degraded and time-accurate computations became very expensive. The current geometry for the LOX boost turbopump has various rotating and stationary components, such as inducer, stators, kicker, hydrolic turbine, where the flow is extremely unsteady. Figure 1 shows the geometry and computed surface pressure of the inducer. The inducer and the hydrolic turbine rotate in different rotational speed.

Development of the Multiple Use Plug Hybrid for Nanosats (MUPHyN) miniature thruster

NASA Astrophysics Data System (ADS)

Eilers, Shannon

The Multiple Use Plug Hybrid for Nanosats (MUPHyN) prototype thruster incorporates solutions to several major challenges that have traditionally limited the deployment of chemical propulsion systems on small spacecraft. The MUPHyN thruster offers several features that are uniquely suited for small satellite applications. These features include 1) a non-explosive ignition system, 2) non-mechanical thrust vectoring using secondary fluid injection on an aerospike nozzle cooled with the oxidizer flow, 3) a non-toxic, chemically-stable combination of liquid and inert solid propellants, 4) a compact form factor enabled by the direct digital manufacture of the inert solid fuel grain. Hybrid rocket motors provide significant safety and reliability advantages over both solid composite and liquid propulsion systems; however, hybrid motors have found only limited use on operational vehicles due to 1) difficulty in modeling the fuel flow rate 2) poor volumetric efficiency and/or form factor 3) significantly lower fuel flow rates than solid rocket motors 4) difficulty in obtaining high combustion efficiencies. The features of the MUPHyN thruster are designed to offset and/or overcome these shortcomings. The MUPHyN motor design represents a convergence of technologies, including hybrid rocket regression rate modeling, aerospike secondary injection thrust vectoring, multiphase injector modeling, non-pyrotechnic ignition, and nitrous oxide regenerative cooling that address the traditional challenges that limit the use of hybrid rocket motors and aerospike nozzles. This synthesis of technologies is unique to the MUPHyN thruster design and no comparable work has been published in the open literature.

Thermophysical Property Testing Using Transient Techniques.

DTIC Science & Technology

1984-06-29

WORDS (Continue on reverse side if necessary and identify by block number) Specific heat HMX carbon/carbon Diffusivity RDX solid propellants Conductivity...energetic materials (AP, " HMX , RDX and HTPB) used in solid rocket fuel to carbon/carbon materials used as rocket nozzles. Studies on AP included single...32 4.1b HMX and RDX ............................35 a 4.2 Carbon/Carbon Materials ...................... 36 5.0 SUMMARY

Estimation of ICBM (Intercontinental Ballistic Missile) Performance Parameters

DTIC Science & Technology

1985-12-01

Formulation . . . . . 42 Staging Event Detection . . . . . . 43 Staging Estimator for Two State System . 46 * Staging Time and Vehicle Parameter...6 4. Land Based Sensor Coordinate System . . . . 10 5. Radar Site Geometry . . . . . . . . . 1 6. Observation Geometry . . . . . . . . . 12 7...Ve dm mi + dmi+d Figure 3. Rocket Thrust of fuel, the equation of motion of the rocket can be devel--S oped. This is a closed system of particles

Los Alamos Novel Rocket Design Flight Tested

ScienceCinema

Tappan, Bryce

2018-04-16

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

Los Alamos Novel Rocket Design Flight Tested

DOE Office of Scientific and Technical Information (OSTI.GOV)

Tappan, Bryce

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

High-Temperature Polymer Composites Tested for Hypersonic Rocket Combustor Backup Structure

NASA Technical Reports Server (NTRS)

Sutter, James K.; Shin, E. Eugene; Thesken, John C.; Fink, Jeffrey E.

2005-01-01

Significant component weight reductions are required to achieve the aggressive thrust-toweight goals for the Rocket Based Combined Cycle (RBCC) third-generation, reusable liquid propellant rocket engine, which is one possible engine for a future single-stage-toorbit vehicle. A collaboration between the NASA Glenn Research Center and Boeing Rocketdyne was formed under the Higher Operating Temperature Propulsion Components (HOTPC) program and, currently, the Ultra-Efficient Engine Technology (UEET) Project to develop carbon-fiber-reinforced high-temperature polymer matrix composites (HTPMCs). This program focused primarily on the combustor backup structure to replace all metallic support components with a much lighter polymer-matrixcomposite- (PMC-) titanium honeycomb sandwich structure.

Diffuse interfacelets in transcritical flows of propellants into high-pressure combustors

NASA Astrophysics Data System (ADS)

Urzay, Javier; Jofre, Lluis

2017-11-01

Rocket engines and new generations of high-power jet engines and diesel engines oftentimes involve the injection of one or more reactants at subcritical temperatures into combustor environments at high pressures, and more particularly, at pressures higher than those corresponding to the critical points of the individual components of the mixture, which typically range from 13 to 50 bars for most propellants. This class of trajectories in the thermodynamic space has been traditionally referred to as transcritical. Under particular conditions often found in hydrocarbon-fueled chemical propulsion systems, and despite the prevailing high pressures, the flow in the combustor may contain regions close to the injector where a diffuse interface is formed in between the fuel and oxidizer streams that is sustained by surface-tension forces as a result of the elevation of the critical pressure of the mixture. This talk describes progress towards modeling these effects in the conservation equations. Funded by the US Department of Energy.

On use of hybrid rocket propulsion for suborbital vehicles

NASA Astrophysics Data System (ADS)

Okninski, Adam

2018-04-01

While the majority of operating suborbital rockets use solid rocket propulsion, recent advancements in the field of hybrid rocket motors lead to renewed interest in their use in sounding rockets. This paper presents results of optimisation of sounding rockets using hybrid propulsion. An overview of vehicles under development during the last decade, as well as heritage systems is provided. Different propellant combinations are discussed and their performance assessment is given. While Liquid Oxygen, Nitrous Oxide and Nitric Acid have been widely tested with various solid fuels in flight, Hydrogen Peroxide remains an oxidiser with very limited sounding rocket applications. The benefits of hybrid propulsion for sounding rockets are given. In case of hybrid rocket motors the thrust curve can be optimised for each flight, using a flow regulator, depending on the payload and mission. Results of studies concerning the optimal burn duration and nozzle selection are given. Specific considerations are provided for the Polish ILR-33 "Amber" sounding rocket. Low regression rates, which up to date were viewed as a drawback of hybrid propulsion may be used to the benefit of maximising rocket performance if small solid rocket boosters are used during the initial flight period. While increased interest in hybrid propulsion is present, no up-to-date reference concerning use of hybrid rocket propulsion for sounding rockets is available. The ultimate goal of the paper is to provide insight into the sensitivity of different design parameters on performance of hybrid sounding rockets and delve into the potential and challenges of using hybrid rocket technology for expendable suborbital applications.

Design Considerations for Clean QED Fusion Propulsion Systems

NASA Astrophysics Data System (ADS)

Bussard, Robert W.; Jameson, Lorin W.

1994-07-01

The direct production of electric power appears possible from fusion reactions between fuels whose products consist solely of charged particles and thus do not present radiation hazards from energetic neutron production, as do reactions involving deuteron-bearing fuels. Among these are the fuels p, 11B, 3He, and 6Li. All of these can be ``burned'' in inertial-electrostatic-fusion (IEF) devices to power QED fusion-electric rocket engines. These IEF sources provide direct-converted electrical power at high voltage (MeV) to drive e-beams for efficient propellant heating to extreme temperatures, with resulting high specific impulse performance capabilities. IEF/QED engine systems using p11B can outperform all other advanced concepts for controlled fusion propulsion by 2-3 orders of magnitude, while 6Li6Li fusion yields one order of magnitude less advance. Either of these fusion rocket propulsion systems can provide very rapid transit for solar system missions, with high payload fractions in single-stage vehicles. The 3He3He reaction can not be used practically for direct electric conversion because of the wide spread in energy of its fusion products. However, it may eventually prove useful for thermal/electrical power generation in central station power plants, or for direct-fusion-product (DFP) propellant heatingin advanced deep-space rocket engines.

Space shuttle solid rocket booster recovery system definition, volume 1

NASA Technical Reports Server (NTRS)

1973-01-01

The performance requirements, preliminary designs, and development program plans for an airborne recovery system for the space shuttle solid rocket booster are discussed. The analyses performed during the study phase of the program are presented. The basic considerations which established the system configuration are defined. A Monte Carlo statistical technique using random sampling of the probability distribution for the critical water impact parameters was used to determine the failure probability of each solid rocket booster component as functions of impact velocity and component strength capability.

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.

Effect of Crystal Orientation on Fatigue Failure of Single Crystal Nickel Base Turbine Blade Superalloys

NASA Technical Reports Server (NTRS)

Arakere, Nagaraj K.; Swanson, Gregory R.

2000-01-01

High Cycle Fatigue (HCF) induced failures in aircraft gas-turbine engines is a pervasive problem affecting a wide range of components and materials. HCF is currently the primary cause of component failures in gas turbine aircraft engines. Turbine blades in high performance aircraft and rocket engines are increasingly being made of single crystal nickel superalloys. Single-crystal Nickel-base superalloys were developed to provide superior creep, stress rupture, melt resistance and thermomechanical fatigue capabilities over polycrystalline alloys previously used in the production of turbine blades and vanes. Currently the most widely used single crystal turbine blade superalloys are PWA 1480/1493 and PWA 1484. These alloys play an important role in commercial, military and space propulsion systems. PWA1493, identical to PWA1480, but with tighter chemical constituent control, is used in the NASA SSME (Space Shuttle Main Engine) alternate turbopump, a liquid hydrogen fueled rocket engine. Objectives for this paper are motivated by the need for developing failure criteria and fatigue life evaluation procedures for high temperature single crystal components, using available fatigue data and finite element modeling of turbine blades. Using the FE (finite element) stress analysis results and the fatigue life relations developed, the effect of variation of primary and secondary crystal orientations on life is determined, at critical blade locations. The most advantageous crystal orientation for a given blade design is determined. Results presented demonstrates that control of secondary and primary crystallographic orientation has the potential to optimize blade design by increasing its resistance to fatigue crack growth without adding additional weight or cost.

Turbopump systems for liquid rocket engines

NASA Technical Reports Server (NTRS)

1974-01-01

The turbopump system, from preliminary design through rocket engine testing is examined. Selection of proper system type for each application and integration of the components into a working system are dealt with. Details are also given on the design of various components including inducers, pumps, turbines, gears, and bearings.

Potential Climate and Ozone Impacts From Hybrid Rocket Engine Emissions

NASA Astrophysics Data System (ADS)

Ross, M.

2009-12-01

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

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

NASA Technical Reports Server (NTRS)

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

1960-01-01

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

Compact Fuel Element Environment Test

NASA Technical Reports Server (NTRS)

Bradley, D. E.; Mireles, O. R.; Hickman, R. R.; Broadway, J. W.

2012-01-01

Deep space missions with large payloads require high specific impulse (I(sub sp)) and relatively high thrust to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average I(sub sp). Nuclear thermal rockets (NTRs) capable of high I(sub sp) thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen, which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3,000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements that employ high melting point metals, ceramics, or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high-temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via noncontact radio frequency heating and expose samples to hydrogen for typical mission durations has been developed to assist in optimal material and manufacturing process selection without employing fissile material. This Technical Memorandum details the test bed design and results of testing conducted to date.

Saturn Apollo Program

NASA Image and Video Library

1961-05-16

On October 27, 1961, the Marshall Space Flight Center (MSFC) and the Nation marked a high point in the 3-year-old Saturn development program when the first Saturn vehicle flew a flawless 215-mile ballistic trajectory from Cape Canaveral, Florida. SA-1 is pictured here, five months before launch, in the MSFC test stand on May 16, 1961. Developed and tested at MSFC under the direction of Dr. Wernher von Braun, SA-1 incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet. and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks, as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle.

Around Marshall

NASA Image and Video Library

2006-07-14

A model of the new Aries I crew launch vehicle, for which NASA is designing, testing and evaluating hardware and related systems, is seen here on display at the Marshall Space Fight Center (MSFC), in Huntsville, Alabama. The Ares I crew launch vehicle is the rocket that will carry a new generation of space explorers into orbit. Under the goals of the Vision for Space Exploration, Ares I is a chief component of the cost-effective space transportation infrastructure being developed by NASA’s Constellation Program. These transportation systems will safely and reliably carry human explorers back to the moon, and then onward to Mars and other destinations in the solar system. The Ares I effort includes multiple project element teams at NASA centers and contract organizations around the nation, and is led by the Exploration Launch Projects Office at NASA’s MFSC. Together, these teams are developing vehicle hardware, evolving proven technologies, and testing components and systems. Their work builds on powerful, reliable space shuttle propulsion elements and nearly a half-century of NASA space flight experience and technological advances. Ares I is an inline, two-stage rocket configuration topped by the Crew Exploration Vehicle, its service module and a launch abort system. The launch vehicle’s first stage is a single, five-segment reusable solid rocket booster derived from the Space Shuttle Program’s reusable solid rocket motor that burns a specially formulated and shaped solid propellant called polybutadiene acrylonitrile (PBAN). The second or upper stage will be propelled by a J-2X main engine fueled with liquid oxygen and liquid hydrogen. In addition to its primary mission of carrying crews of four to six astronauts to Earth orbit, the launch vehicle’s 25-ton payload capacity might be used for delivering cargo to space, bringing resources and supplies to the International Space Station or dropping payloads off in orbit for retrieval and transport to exploration teams on the moon. Crew transportation to the space station is planned to begin no later than 2014. The first lunar excursion is scheduled for the 2020 timeframe.

Rocket Launch-Induced Vibration and Ignition Overpressure Response

NASA Technical Reports Server (NTRS)

Caimi, Raoul; Margasahayam, Ravi; Nayfeh, Jamal; Thompson, Karen (Technical Monitor)

2001-01-01

Rocket-induced vibration and ignition overpressure response environments are predicted in the low-frequency (5 to 200 hertz) range. The predictions are necessary to evaluate their impact on critical components, structures, and facilities in the immediate vicinity of the rocket launch pad.

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.

Evaluation of on-board hydrogen storage methods for hypersonic vehicles

NASA Technical Reports Server (NTRS)

Akyurtlu, Ates; Akyurtlu, J. F.; Adeyiga, A. A.; Perdue, Samara; Northam, G. B.

1989-01-01

Hydrogen is the foremost candidate as a fuel for use in high speed transport. Since any aircraft moving at hypersonic speeds must have a very slender body, means of decreasing the storage volume requirements below that for liquid hydrogen are needed. The total performance of the hypersonic plane needs to be considered for the evaluation of candidate fuel and storage systems. To accomplish this, a simple model for the performance of a hypersonic plane is presented. To allow for the use of different engines and fuels during different phases of flight, the total trajectory is divided into three phases: subsonic-supersonic, hypersonic and rocket propulsion phase. The fuel fraction for the first phase is found be a simple energy balance using an average thrust to drag ratio for this phase. The hypersonic flight phase is investigated in more detail by taking small altitude increments. This approach allowed the use of flight profiles other than the constant dynamic pressure flight. The effect of fuel volume on drag, structural mass and tankage mass was introduced through simplified equations involving the characteristic dimension of the plane. The propellant requirement for the last phase is found by employing the basic rocket equations. The candidate fuel systems such as the cryogenic fuel combinations and solid and liquid endothermic hydrogen generators are first screened thermodynamically with respect to their energy densities and cooling capacities and then evaluated using the above model.

An innovative demonstration of high power density in a compact MDH (magnetohydrodynamic) generator

NASA Astrophysics Data System (ADS)

Schmidt, H. J.; Lineberry, J. T.; Chapman, J. N.

1990-06-01

The present program was conducted by the University of Tennessee Space Institute (UTSI). It was by its nature a high risk experimental program to demonstrate the feasibility of high power density operation in a laboratory scale combustion driven MHD generator. Maximization of specific energy was not a consideration for the present program, but the results have implications in this regard by virtue of high energy fuel used. The power density is the ratio of the electrical energy output to the internal volume of the generator channel. The MHD process is a volumetric process and the power density is therefore a direct measure of the compactness of the system. Specific energy, is the ratio of the electrical energy output to consumable energy used for its production. The two parameters are conceptually interrelated. To achieve high power density and implied commensurate low system volume and weight, it was necessary to use an energetic fuel. The high energy fuel of choice was a mixture of powdered aluminum and carbon seeded with potassium carbonate and burned with gaseous oxygen. The solid fuel was burned in a hybrid combustion scheme wherein the fuel was cast within a cylindrical combustor in analogy with a solid propellant rocket motor. Experimental data is limited to gross channel output current and voltage, magnetic field strength, fuel and oxidizer flow rates, flow train external temperatures and combustor pressure. Similarly, while instantaneous oxidizer flow rates were measured, only average fuel consumption based on pre and post test component weights and dimensions was possible.

Transient Modeling of Hybrid Rocket Low Frequency Instabilities

NASA Technical Reports Server (NTRS)

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

2003-01-01

A comprehensive dynamic model of a hybrid rocket has been developed in order to understand and predict the transient behavior including instabilities. A linearized version of the transient model predicted the low-frequency chamber pressure oscillations that are commonly observed in hybrids. The source of the instabilities is based on a complex coupling of thermal transients in the solid fuel, wall heat transfer blocking due to fuel regression rate and the transients in the boundary layer that forms on the fuel surface. The oscillation frequencies predicted by the linearized theory are in very good agreement with 43 motor test results obtained from the hybrid propulsion literature. The motor test results used in the comparison cover a very wide spectrum of parameters including: 1) four separate research and development programs, 2) three different oxidizers (LOX, GOX, N2O), 3) a wide range of motor dimensions (i.e. from 5 inch diameter to 72 inch diameter) and operating conditions and 4) several fuel formulations. A simple universal scaling formula for the frequency of the primary oscillation mode is suggested.

Studies of the thermal and optical responses of H atoms in solid H2

NASA Technical Reports Server (NTRS)

Gaines, James R.; Vause, Chester A., III

1990-01-01

It was the goal of this reserch project to model both the storage of energy in solid hydrogen in the form of atoms and the conversion of this stored energy into other forms of useful energy. The basic ideas of rocket propulsion originate in classical physics and they remain unchanged. To escape a strong gravitational field, the 'burn time' must be minimized but in negligible force fields, the burn time is unimportant and only the relative masses of rocket to fuel determine a specific exhaust velocity. It is in this latter case that low mass fuels such as hydrogen become very important. The burning of hydrogen in oxygen is a 'benchmark' fuel today providing a specific impulse of 400 seconds or better. More exotic fuels will be needed for many of the interesting explorations of the future but they still must have large energy releases per unit mass. It is in this context that propulsion based on hydrogen atom recombination receives attention and these studies will serve as engineering guides.

Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Anderson, William E.

2005-01-01

The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

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

NASA Technical Reports Server (NTRS)

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

2003-01-01

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

Rocket-Based Combined Cycle Engine Concept Development

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

Fundamental rocket injector/spray programs at the Phillips Laboratory

NASA Astrophysics Data System (ADS)

Talley, D. G.

1993-11-01

The performance and stability of liquid rocket engines is determined to a large degree by atomization, mixing, and combustion processes. Control over these processes is exerted through the design of the injector. Injectors in liquid rocket engines are called upon to perform many functions. They must first of all mix the propellants to provide suitable performance in the shortest possible length. For main injectors, this is driven by the tradeoff between the combustion chamber performance, stability, efficiency, and its weight and cost. In gas generators and preburners, however, it is also driven by the possibility of damage to downstream components, for example piping and turbine blades. This can occur if unburned fuel and oxidant later react to create hot spots. Weight and cost considerations require that the injector design be simple and lightweight. For reusable engines, the injectors must also be durable and easily maintained. Suitable atomization and mixing must be produced with as small a pressure drop as possible, so that the size and weight of pressure vessels and turbomachinery can be minimized. However, the pressure drop must not be so small as to promote feed system coupled instabilities. Another important function of the injectors is to ensure that the injector face plate and the chamber and nozzle walls are not damaged. Typically this requires reducing the heat transfer to an acceptable level and also keeping unburned oxygen from chemically attacking the walls, particularly in reusable engines. Therefore the mixing distribution is often tailored to be fuel-rich near the walls. Wall heat transfer can become catastrophically damaging in the presence of acoustic instabilities, so the injector must prevent these from occurring at all costs. In addition to acoustic stability (but coupled with it), injectors must also be kinetically stable. That is, the flame itself must maintain ignition in the combustion chamber. This is not typically a problem with main injectors, but can be a consideration in preburners, where the desire to keep turbine inlet temperatures as cool as possible can make it advantageous for the preburners to operate as far from stoichiometry as can be tolerated.

The composition-explicit distillation curve technique: Relating chemical analysis and physical properties of complex fluids.

PubMed

Bruno, Thomas J; Ott, Lisa S; Lovestead, Tara M; Huber, Marcia L

2010-04-16

The analysis of complex fluids such as crude oils, fuels, vegetable oils and mixed waste streams poses significant challenges arising primarily from the multiplicity of components, the different properties of the components (polarity, polarizability, etc.) and matrix properties. We have recently introduced an analytical strategy that simplifies many of these analyses, and provides the added potential of linking compositional information with physical property information. This aspect can be used to facilitate equation of state development for the complex fluids. In addition to chemical characterization, the approach provides the ability to calculate thermodynamic properties for such complex heterogeneous streams. The technique is based on the advanced distillation curve (ADC) metrology, which separates a complex fluid by distillation into fractions that are sampled, and for which thermodynamically consistent temperatures are measured at atmospheric pressure. The collected sample fractions can be analyzed by any method that is appropriate. The analytical methods we have applied include gas chromatography (with flame ionization, mass spectrometric and sulfur chemiluminescence detection), thin layer chromatography, FTIR, corrosivity analysis, neutron activation analysis and cold neutron prompt gamma activation analysis. By far, the most widely used analytical technique we have used with the ADC is gas chromatography. This has enabled us to study finished fuels (gasoline, diesel fuels, aviation fuels, rocket propellants), crude oils (including a crude oil made from swine manure) and waste oils streams (used automotive and transformer oils). In this special issue of the Journal of Chromatography, specifically dedicated to extraction technologies, we describe the essential features of the advanced distillation curve metrology as an analytical strategy for complex fluids. Published by Elsevier B.V.

Ames Hybrid Combustion Facility

NASA Technical Reports Server (NTRS)

Zilliac, Greg; Karabeyoglu, Mustafa A.; Cantwell, Brian; Hunt, Rusty; DeZilwa, Shane; Shoffstall, Mike; Soderman, Paul T.; Bencze, Daniel P. (Technical Monitor)

2003-01-01

The report summarizes the design, fabrication, safety features, environmental impact, and operation of the Ames Hybrid-Fuel Combustion Facility (HCF). The facility is used in conducting research into the scalability and combustion processes of advanced paraffin-based hybrid fuels for the purpose of assessing their applicability to practical rocket systems. The facility was designed to deliver gaseous oxygen at rates between 0.5 and 16.0 kg/sec to a combustion chamber operating at pressures ranging from 300 to 900. The required run times were of the order of 10 to 20 sec. The facility proved to be robust and reliable and has been used to generate a database of regression-rate measurements of paraffin at oxygen mass flux levels comparable to those of moderate-sized hybrid rocket motors.

Performance Comparison of Axisymmetric and Three-dimensional Hydrogen Film Coolant Injection in a 110N Hydrogen/oxygen Rocket

NASA Technical Reports Server (NTRS)

Arrington, Lynn A.; Reed, Brian D.

1992-01-01

An experimental performance comparison of two geometrically different fuel film coolant injection sleeves was conducted on a 110 N gaseous hydrogen/oxygen rocket. One sleeve had slots milled axially down the walls and the other had a smooth surface to give axisymmetric flow. The comparison was made to investigate a conclusion in an earlier study that attributed a performance underprediction to a symplifying modeling assumption of axisymmetric fuel film flow. The smooth sleeve had higher overall performance at one film coolant percentage and approximately the same or slightly better at another. The study showed that the lack of modeling of three-dimensional effects was not the cause of the performance underprediction as speculated in earlier analytical studies.

Internship at NASA Kennedy Space Center's Cryogenic Test laboratory

NASA Technical Reports Server (NTRS)

Holland, Katherine

2013-01-01

NASA's Kennedy Space Center (KSC) is known for hosting all of the United States manned rocket launches as well as many unmanned launches at low inclinations. Even though the Space Shuttle recently retired, they are continuing to support unmanned launches and modifying manned launch facilities. Before a rocket can be launched, it has to go through months of preparation, called processing. Pieces of a rocket and its payload may come in from anywhere in the nation or even the world. The facilities all around the center help integrate the rocket and prepare it for launch. As NASA prepares for the Space Launch System, a rocket designed to take astronauts beyond Low Earth Orbit throughout the solar system, technology development is crucial for enhancing launch capabilities at the KSC. The Cryogenics Test Laboratory at Kennedy Space Center greatly contributes to cryogenic research and technology development. The engineers and technicians that work there come up with new ways to efficiently store and transfer liquid cryogens. NASA has a great need for this research and technology development as it deals with cryogenic liquid hydrogen and liquid oxygen for rocket fuel, as well as long term space flight applications. Additionally, in this new era of space exploration, the Cryogenics Test Laboratory works with the commercial sector. One technology development project is the Liquid Hydrogen (LH2) Ground Operations Demonstration Unit (GODU). LH2 GODU intends to demonstrate increased efficiency in storing and transferring liquid hydrogen during processing, loading, launch and spaceflight of a spacecraft. During the Shuttle Program, only 55% of hydrogen purchased was used by the Space Shuttle Main Engines. GODU's goal is to demonstrate that this percentage can be increased to 75%. Figure 2 shows the GODU layout when I concluded my internship. The site will include a 33,000 gallon hydrogen tank (shown in cyan) with a heat exchanger inside the hydrogen tank attached to a refrigerator capable of removing 850 Watts at 20 Kelvin (shown in green). The refrigerator and most of its supporting equipment will be kept in a standard shipping container (shown in pink). Currently, GODU is in the fabrication process and some of the large components have already been purchased.

A Review of Gas-Cooled Reactor Concepts for SDI Applications

DTIC Science & Technology

1989-08-01

710 program .) Wire- Core Reactor (proposed by Rockwell). The wire- core reactor utilizes thin fuel wires woven between spacer wires to form an open...reactor is based on results of developmental studies of nuclear rocket propulsion systems. The reactor core is made up of annular fuel assemblies of...XE Addendum to Volume II. NERVA Fuel Development , Westinghouse Astronuclear Laboratory, TNR-230, July 15’ 1972. J I8- Rover Program Reactor Tests

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

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

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

27 CFR 447.21 - The U.S. Munitions Import List.

Code of Federal Regulations, 2010 CFR

2010-04-01

...—launch vehicles, guided missiles, ballistic missiles, rockets, torpedoes, bombs and mines (a) Rockets (including but not limited to meteorological and other sounding rockets), bombs, grenades, torpedoes, depth..., the following: Fuses and components for the items in this category, bomb racks and shackles, bomb...

Orbital transfer rocket engine technology 7.5K-LB thrust rocket engine preliminary design

NASA Technical Reports Server (NTRS)

Harmon, T. J.; Roschak, E.

1993-01-01

A preliminary design of an advanced LOX/LH2 expander cycle rocket engine producing 7,500 lbf thrust for Orbital Transfer vehicle missions was completed. Engine system, component and turbomachinery analysis at both on design and off design conditions were completed. The preliminary design analysis results showed engine requirements and performance goals were met. Computer models are described and model outputs are presented. Engine system assembly layouts, component layouts and valve and control system analysis are presented. Major design technologies were identified and remaining issues and concerns were listed.

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.

Parametric study and performance analysis of hybrid rocket motors with double-tube configuration

NASA Astrophysics Data System (ADS)

Yu, Nanjia; Zhao, Bo; Lorente, Arnau Pons; Wang, Jue

2017-03-01

The practical implementation of hybrid rocket motors has historically been hampered by the slow regression rate of the solid fuel. In recent years, the research on advanced injector designs has achieved notable results in the enhancement of the regression rate and combustion efficiency of hybrid rockets. Following this path, this work studies a new configuration called double-tube characterized by injecting the gaseous oxidizer through a head end injector and an inner tube with injector holes distributed along the motor longitudinal axis. This design has demonstrated a significant potential for improving the performance of hybrid rockets by means of a better mixing of the species achieved through a customized injection of the oxidizer. Indeed, the CFD analysis of the double-tube configuration has revealed that this design may increase the regression rate over 50% with respect to the same motor with a conventional axial showerhead injector. However, in order to fully exploit the advantages of the double-tube concept, it is necessary to acquire a deeper understanding of the influence of the different design parameters in the overall performance. In this way, a parametric study is carried out taking into account the variation of the oxidizer mass flux rate, the ratio of oxidizer mass flow rate injected through the inner tube to the total oxidizer mass flow rate, and injection angle. The data for the analysis have been gathered from a large series of three-dimensional numerical simulations that considered the changes in the design parameters. The propellant combination adopted consists of gaseous oxygen as oxidizer and high-density polyethylene as solid fuel. Furthermore, the numerical model comprises Navier-Stokes equations, k-ε turbulence model, eddy-dissipation combustion model and solid-fuel pyrolysis, which is computed through user-defined functions. This numerical model was previously validated by analyzing the computational and experimental results obtained for conventional hybrid rocket designs. In addition, a performance analysis is conducted in order to evaluate the influence in the performance provoked by the possible growth of the diameter of the inner fuel grain holes during the motor operation. The latter phenomenon is known as burn through holes. Finally, after a statistical analysis of the data, a regression rate expression as a function of the design parameters is obtained.

Status Update Report for the Peregrine 100km Sounding Rocket Project

NASA Technical Reports Server (NTRS)

Dyer, Jonny; Zilliac, Greg; Doran, Eric; Marzona, Mark Thadeus; Lohner, Kevin; Karlik, Evan; Cantwell, Brian; Karabeyoglu, Arif

2008-01-01

The Peregrine Sounding Rocket Program is a joint basic research program of NASA Ames Research Center, NASA Wallops, Stanford University and the Space Propulsion Group, Inc. (SPG). The goal is to determine the applicability of liquifying hybrid technology to a small launch system. The approach is to design, build, test and y a stable, efficient liquefying fuel hybrid rocket vehicle to an altitude of 100 km. The program was kicked o in October of 2006 and has seen considerable progress in the subsequent 18 months. Two virtually identical vehicles will be constructed and own out of the NASA Sounding Rocket Facility at Wallops Island. This paper presents the current status of the project as of June 2008. For background on the project, the reader is referred to last year's paper.

KSC-2009-5819

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – Tugboats safely deliver the Pegasus barge, carrying external tank 134, to the dock in the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5818

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – Tugboats safely deliver the Pegasus barge, carrying external tank 134, to the dock in the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5825

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – External tank 134 is towed toward the open door of the transfer aisle of the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. The Pegasus barge, carrying the fuel tank, arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. Next, the tank will be transported into the VAB where it will be stored until needed. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5821

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – A worker tows external tank 134 off the Pegasus barge docked in the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After the fuel tank is offloaded, it will be transported into the VAB where it will be stored until needed. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5817

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – Tugboats safely deliver the Pegasus barge, carrying external tank 134, to the dock in the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5820

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – Workers prepare to offload external tank 134 from the Pegasus barge docked in the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5824

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – External tank 134 is towed toward the looming 525-foot-tall Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. The Pegasus barge, carrying the fuel tank, arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. Next, the tank will be transported into the VAB where it will be stored until needed. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

Turbopump Design and Analysis Approach for Nuclear Thermal Rockets

NASA Technical Reports Server (NTRS)

Chen, Shu-cheng S.; Veres, Joseph P.; Fittje, James E.

2006-01-01

A rocket propulsion system, whether it is a chemical rocket or a nuclear thermal rocket, is fairly complex in detail but rather simple in principle. Among all the interacting parts, three components stand out: they are pumps and turbines (turbopumps), and the thrust chamber. To obtain an understanding of the overall rocket propulsion system characteristics, one starts from analyzing the interactions among these three components. It is therefore of utmost importance to be able to satisfactorily characterize the turbopump, level by level, at all phases of a vehicle design cycle. Here at NASA Glenn Research Center, as the starting phase of a rocket engine design, specifically a Nuclear Thermal Rocket Engine design, we adopted the approach of using a high level system cycle analysis code (NESS) to obtain an initial analysis of the operational characteristics of a turbopump required in the propulsion system. A set of turbopump design codes (PumpDes and TurbDes) were then executed to obtain sizing and performance characteristics of the turbopump that were consistent with the mission requirements. A set of turbopump analyses codes (PUMPA and TURBA) were applied to obtain the full performance map for each of the turbopump components; a two dimensional layout of the turbopump based on these mean line analyses was also generated. Adequacy of the turbopump conceptual design will later be determined by further analyses and evaluation. In this paper, descriptions and discussions of the aforementioned approach are provided and future outlooks are discussed.

Characterisation of oxidation products of 1,1-dimethylhydrazine by high-resolution orbitrap mass spectrometry.

PubMed

Ul'yanovskii, N V; Kosyakov, D S; Pikovskoi, I I; Khabarov, Yu G

2017-05-01

1,1-Dimethylhydrazine is used as a fuel for carrier rockets in the majority of countries implementing space exploration programs. Being highly reactive, 1,1-dimethylhydrazine easily undergoes oxidative transformation with the formation of a number of toxic, mutagenic, and teratogenic compounds. The use of high-resolution mass spectrometry for the study of the reaction of 1,1-dimethylhydrazine oxidation with hydrogen peroxide in aqueous solution allowed us to find hundreds of nitrogen-containing products of the CHN and CHNO classes, formed via radical processes. The vast majority of the compounds have not been previously considered as possible products of the transformation of rocket fuel. We have shown that the oxidation of 1,1-dimethylhydrazine proceeds in two stages, with the formation of a great number of complex unstable intermediates that contain up to ten nitrogen atoms. These intermediates are subsequently converted into final reaction products with a concomitant decrease in the average molecular weight. The intermediates and final products of the oxidative transformation of 1,1-dimethylhydrazine were characterised on the basis of their elemental composition using van Krevelen diagrams and possible compounds corresponding to the most intense peaks in the mass spectra were proposed. The data obtained are indicative of the presence of the following classes of heterocyclic nitrogen-containing compounds among the oxidation products: imines, piperidines, pyrrolidines, dihydropyrazoles, dihydroimidazoles, triazoles, aminotriazines, and tetrazines. The results obtained open up possibilities for the targeted search and identification of new toxic products of the degradation of rocket fuel and, as a result, a more adequate assessment of the ecological consequences of space-rocket activity. Copyright © 2017 Elsevier Ltd. All rights reserved.

Modeling Transients and Designing a Passive Safety System for a Nuclear Thermal Rocket Using Relap5

NASA Astrophysics Data System (ADS)

Khatry, Jivan

Long-term high payload missions necessitate the need for nuclear space propulsion. Several nuclear reactor types were investigated by the Nuclear Engine for Rocket Vehicle Application (NERVA) program of National Aeronautics and Space Administration (NASA). Study of planned/unplanned transients on nuclear thermal rockets is important due to the need for long-term missions. A NERVA design known as the Pewee I was selected for this purpose. The following transients were run: (i) modeling of corrosion-induced blockages on the peripheral fuel element coolant channels and their impact on radiation heat transfer in the core, and (ii) modeling of loss-of-flow-accidents (LOFAs) and their impact on radiation heat transfer in the core. For part (i), the radiation heat transfer rate of blocked channels increases while their neighbors' decreases. For part (ii), the core radiation heat transfer rate increases while the flow rate through the rocket system is decreased. However, the radiation heat transfer decreased while there was a complete LOFA. In this situation, the peripheral fuel element coolant channels handle the majority of the radiation heat transfer. Recognizing the LOFA as the most severe design basis accident, a passive safety system was designed in order to respond to such a transient. This design utilizes the already existing tie rod tubes and connects them to a radiator in a closed loop. Hence, this is basically a secondary loop. The size of the core is unchanged. During normal steady-state operation, this secondary loop keeps the moderator cool. Results show that the safety system is able to remove the decay heat and prevent the fuel elements from melting, in response to a LOFA and subsequent SCRAM.

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

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

16 CFR § 1500.85 - Exemptions from classification as banned hazardous substances.

Code of Federal Regulations, 2013 CFR

2013-01-01

... component has no hazards other than being in a self-pressurized container. (8) Model rocket propellant devices designed for use in light-weight, recoverable, and reflyable model rockets, provided such devices... recovery system activation devices intended for use with premanufactured model rocket engines wherein all...

16 CFR 1500.85 - Exemptions from classification as banned hazardous substances.

Code of Federal Regulations, 2014 CFR

2014-01-01

... component has no hazards other than being in a self-pressurized container. (8) Model rocket propellant devices designed for use in light-weight, recoverable, and reflyable model rockets, provided such devices... recovery system activation devices intended for use with premanufactured model rocket engines wherein all...

16 CFR 1500.85 - Exemptions from classification as banned hazardous substances.

Code of Federal Regulations, 2011 CFR

2011-01-01

... component has no hazards other than being in a self-pressurized container. (8) Model rocket propellant devices designed for use in light-weight, recoverable, and reflyable model rockets, provided such devices... recovery system activation devices intended for use with premanufactured model rocket engines wherein all...

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.

Optical Imaging of Flow Pattern and Phantom

NASA Technical Reports Server (NTRS)

Galland, Pierre A.; Liang, X.; Wang, L.; Ho, P. P.; Alfano, R. R.; Breisacher, K.

1999-01-01

Time-resolved optical imaging technique has been used to image the spatial distribution of small droplets and jet sprays in a highly scattering environment. The snake and ballistic components of the transmitted pulse are less scattered, and contain direct information about the sample to facilitate image formation as opposed to the diffusive components which are due to multiple collisions as a light pulse propagates through a scattering medium. In a time-gated imaging scheme, these early-arriving, image-bearing components of the incident pulse are selected by opening a gate for an ultrashort period of time and a shadowgram image is detected. Using a single shot cooled CCD camera system, the formation of water droplets is monitored as a function of time. Picosecond time-gated image of drop in scattering cells, spray droplets as a function of let speed and gas pressure, and model calcification samples consisted of calcium carbonate particles of irregular shapes ranging in size from 0. 1 to 1.5 mm affixed to a microscope slide have been measured. Formation produced by an impinging jet will be further monitored using a CCD with 1 kHz framing illuminated with pulsed light. The desired image resolution of the fuel droplets is on the 20 pm scale using early light through a highly scattering medium. A 10(exp -6)m displacement from a jet spray with a flow speed of 100 m/sec introduced by the ns grating pulse used in the imaging is negligible. Early ballistic/snake light imaging offers nondestructive and noninvasive method to observe the spatial distribution of hidden objects inside a highly scattering environment for space, biomedical, and materials applications. In this paper, the techniques we will present are time-resolved K-F transillumination imaging and time-gated scattered light imaging. With a large dynamic range and high resolution, time-gated early light imaging has the potential for improving rocket/aircraft design by determining jets shape and particle sizes. Refinements to these techniques may enable drop size measurements in the highly scattering, optically dense region of multi-element rocket injectors. These types of measurements should greatly enhance the design of stable, and higher performing rocket engines.

Liquid boosters for Shuttle?

NASA Astrophysics Data System (ADS)

Robertson, Donald F.

1989-12-01

The use of liquid rocket boosters (LRBs) for the Space Shuttle is proposed. The advantages LRBs provide are improved flight safety due to the use of four engines instead of two and less environmental pollution than solid rocket boosters because LRBs utilize clean-burning fuels. The LRBs also permit very high launch rates and increased safety in assembly and mating of the Shuttle. Concerns about LRBs such as costs, diameter, support capability, and water recovery are examined.

Propulsion system needs

NASA Technical Reports Server (NTRS)

Gunn, Stanley

1991-01-01

The needs of the designer of a solid core nuclear rocket engine are discussed. Some of the topics covered include: (1) a flight thrust module/feed system module assembly; (2) a nuclear thermal rocket (NTR), expander cycle, dual T/P; (3) turbopump operating conditions; (4) typical system parameters; (5) growth capability composite fuel elements; (6) a NTR radiation cooled nozzle extension; (7) a NFS-3B Feed System; and (8) a NTR Integrated Pneumatic-Fluidics Control System.

41. Historic photo of Building 202 test cell interior, Robert ...

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

41. Historic photo of Building 202 test cell interior, Robert J. Gardener checking fuel implinging qualities of a twenty-thousand-pound-thrust rocket engine injector. Setting appears to be a platform mounted on top of scrubber tank underneath test cell floor, December 1959. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA photo number C-52166. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Battery and Fuel Cell Development for NASA's Constellation Missions

NASA Technical Reports Server (NTRS)

Manzo, Michelle A.

2009-01-01

NASA's return to the moon will require advanced battery, fuel cell and regenerative fuel cell energy storage systems. This paper will provide an overview of the planned energy storage systems for the Orion Spacecraft and the Aries rockets that will be used in the return journey to the Moon. Technology development goals and approaches to provide batteries and fuel cells for the Altair Lunar Lander, the new space suit under development for extravehicular activities (EY A) on the Lunar surface, and the Lunar Surface Systems operations will also be discussed.

Battery and Fuel Cell Development for NASA's Exploration Missions

NASA Technical Reports Server (NTRS)

Manzo, Michelle A.; Reid, Concha M.

2009-01-01

NASA's return to the moon will require advanced battery, fuel cell and regenerative fuel cell energy storage systems. This paper will provide an overview of the planned energy storage systems for the Orion Spacecraft and the Aries rockets that will be used in the return journey to the Moon. Technology development goals and approaches to provide batteries and fuel cells for the Altair Lunar Lander, the new space suit under development for extravehicular activities (EVA) on the Lunar surface, and the Lunar Surface Systems operations will also be discussed.

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.

Nuclear Cryogenic Propulsion Stage Fuel Design and Fabrication

NASA Technical Reports Server (NTRS)

Hickman, Robert; Broadway, Jeramie; Mireles, Omar; Webb, Jon; Qualls, Lou

2012-01-01

Nuclear Cryogenic Propulsion Stage (NCPS) is a game changing technology for space exploration. Goal of assessing the affordability and viability of an NCPS includes these overall tasks: (1) Pre-conceptual design of the NCPS and architecture integration (2) NCPS Fuel Design and Testing (3) Nuclear Thermal Rocket Element Environmental Simulator (NTREES) (4) Affordable NCPS Development and Qualification Strategy (5) Second Generation NCPS Concepts. There is a critical need for fuels development. Fuel task objectives are to demonstrate capabilities and critical technologies using full scale element fabrication and testing.

Nuclear Cryogenic Propulsion Stage Fuel Design and Fabrication

NASA Technical Reports Server (NTRS)

Hickman, Robert; Broadway, Jeramie; Mireles, Omar; Webb, Jon; Qualls, Lou

2012-01-01

Nuclear Cryogenic Propulsion Stage (NCPS) is a game changing technology for space exploration. Goal of assessing the affordability and viability of an NCPS includes thses overall tasks: (1) Pre-conceptual design of the NCPS and architecture integration (2) NCPS Fuel Design and Testing (3) Nuclear Thermal Rocket Element Environmental Simulator (NTREES) (4) Affordable NCPS Development and Qualification Strategy (5) Second Generation NCPS Concepts. There is a critical need for fuels development. Fuel task objectives are to demonstrate capabilities and critical technologies using full scale element fabrication and testing.

Flight and Preflight Tests of a Ram Jet Burning Magnesium Slurry Fuel and Utilizing a Solid-propellant Gas Generator for Fuel Expulsion

NASA Technical Reports Server (NTRS)

Bartlett, Walter, A , jr; Hagginbotham, William K , Jr

1955-01-01

Data obtained from the first flight test of a ram jet utilizing a magnesium slurry fuel are presented. The ram jet accelerated from a Mach number of 1.75 to a Mach number of 3.48 in 15.5 seconds. During this period a maximum values of air specific impulse and gross thrust coefficient were calculated to be 151 seconds and 0.658, respectively. The rocket gas generator used as a fuel-pumping system operated successfully.

Liquid Rocket Engine Testing

DTIC Science & Technology

2016-10-21

Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static-tested in development • This

Nuclear design of a vapor core reactor for space nuclear propulsion

NASA Astrophysics Data System (ADS)

Dugan, Edward T.; Watanabe, Yoichi; Kuras, Stephen A.; Maya, Isaac; Diaz, Nils J.

1993-01-01

Neutronic analysis methodology and results are presented for the nuclear design of a vapor core reactor for space nuclear propulsion. The Nuclear Vapor Thermal Reactor (NVTR) Rocket Engine uses modified NERVA geometry and systems which the solid fuel replaced by uranium tetrafluoride vapor. The NVTR is an intermediate term gas core thermal rocket engine with specific impulse in the range of 1000-1200 seconds; a thrust of 75,000 lbs for a hydrogen flow rate of 30 kg/s; average core exit temperatures of 3100 K to 3400 K; and reactor thermal powers of 1400 to 1800 MW. Initial calculations were performed on epithermal NVTRs using ZrC fuel elements. Studies are now directed at thermal NVTRs that use fuel elements made of C-C composite. The large ZrC-moderated reactors resulted in thrust-to-weight ratios of only 1 to 2; the compact C-C composite systems yield thrust-to-weight ratios of 3 to 5.

High-pressure calorimeter chamber tests for liquid oxygen/kerosene (LOX/RP-1) rocket combustion

NASA Technical Reports Server (NTRS)

Masters, Philip A.; Armstrong, Elizabeth S.; Price, Harold G.

1988-01-01

An experimental program was conducted to investigate the rocket combustion and heat transfer characteristics of liquid oxygen/kerosene (LOX/RP-1) mixtures at high chamber pressures. Two water-cooled calorimeter chambers of different combustion lengths were tested using 37- and 61-element oxidizer-fuel-oxidizer triplet injectors. The tests were conducted at nominal chamber pressures of 4.1, 8.3, and 13.8 MPa abs (600, 1200, and 2000 psia). Heat flux Q/A data were obtained for the entire calorimeter length for oxygen/fuel mixture ratios of 1.8 to 3.3. Test data at 4.1 MPa abs compared favorably with previous test data from another source. Using an injector with a fuel-rich outer zone reduced the throat heat flux by 47 percent with only a 4.5 percent reduction in the characteristic exhaust velocity efficiency C* sub eff. The throat heat transfer coefficient was reduced approximately 40 percent because of carbon deposits on the chamber wall.

Design, Activation, and Operation of the J2-X Subscale Simulator (JSS)

NASA Technical Reports Server (NTRS)

Saunders, Grady P.; Raines, Nickey G.; Varner, Darrel G.

2009-01-01

The purpose of this paper is to give a detailed description of the design, activation, and operation of the J2-X Subscale Simulator (JSS) installed in Cell 1 of the E3 test facility at Stennis Space Center, MS (SSC). The primary purpose of the JSS is to simulate the installation of the J2-X engine in the A3 Subscale Rocket Altitude Test Facility at SSC. The JSS is designed to give aerodynamically and thermodynamically similar plume properties as the J2-X engine currently under development for use as the upper stage engine on the ARES I and ARES V spacecraft. The JSS is a scale pressure fed, LOX/GH fueled rocket that is geometrically similar to the J2-X from the throat to the nozzle exit plane (NEP) and is operated at the same oxidizer to fuel ratios and chamber pressures. This paper describes the heritage hardware used as the basis of the JSS design, the newly designed rocket hardware, igniter systems used, and the activation and operation of the JSS.

Multiple dopant injection system for small rocket engines

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

A preliminary analysis of low frequency pressure oscillations in hybrid rocket motors

NASA Technical Reports Server (NTRS)

Jenkins, Rhonald M.

1994-01-01

Past research with hybrid rockets has suggested that certain motor operating conditions are conducive to the formation of pressure oscillations, or flow instabilities, within the motor combustion chamber. These combustion-related vibrations or pressure oscillations may be encountered in virtually any type of rocket motor and typically fall into three frequency ranges: low frequency oscillations (0-300 Hz); intermediate frequency oscillations (400-1000 Hz); and high frequency oscillations (greater than 1000 Hz). In general, combustion instability is characterized by organized pressure oscillations occurring at well-defined intervals with pressure peaks that may maintain themselves, grow, or die out. Usually, such peaks exceed +/- 5% of the mean chamber pressure. For hybrid motors, these oscillations have been observed to grow to a limiting amplitude which may be dependent on factors such as fuel characteristics, oxidizer injector characteristics, average chamber pressure, oxidizer mass flux, combustion chamber length, and grain geometry. The approach taken in the present analysis is to develop a modified chamber length, L, instability theory which accounts for the relationship between pressure and oxidizer to fuel concentration ratio in the motor.

A fast sampling device for the mass spectrometric analysis of liquid rocket engine exhaust

NASA Technical Reports Server (NTRS)

Ryason, P. R.

1975-01-01

The design of a device to obtain compositional data on rocket exhaust by direct sampling of reactive flow exhausts into a mass spectrometer is presented. Sampling at three stages differing in pressure and orifice angle and diameter is possible. Results of calibration with pure gases and gas mixtures are erratic and of unknown accuracy for H2, limiting the usefulness of the apparatus for determining oxidizer/fuel ratios from combustion product analysis. Deposition effects are discussed, and data obtained from rocket exhaust spectra are analyzed to give O/F ratios and mixture ratio distribution. The O/F ratio determined spectrometrically is insufficiently accurate for quantitative comparison with cold flow data. However, a criterion for operating conditions with improved mixing of fuel and oxidizer which is consistent with cold flow results may be obtained by inspection of contour plots. A chemical inefficiency in the combustion process when oxidizer is in excess is observed from reactive flow measurements. Present results were obtained with N2O4/N2H4 propellants.

A Laboratory Model of a Hydrogen/Oxygen Engine for Combustion and Nozzle Studies

NASA Technical Reports Server (NTRS)

Morren, Sybil Huang; Myers, Roger M.; Benko, Stephen E.; Arrington, Lynn A.; Reed, Brian D.

1993-01-01

A small laboratory diagnostic thruster was developed to augment present low thrust chemical rocket optical and heat flux diagnostics at the NASA Lewis Research Center. The objective of this work was to evaluate approaches for the use of temperature and pressure sensors for the investigation of low thrust rocket flow fields. The nominal engine thrust was 110 N. Tests were performed at chamber pressures of about 255 kPa, 370 kPa, and 500 kPa with oxidizer to fuel mixture ratios between 4.0 and 8.0. Two gaseous hydrogen/gaseous oxygen injector designs were tested with 60 percent and 75 percent fuel film cooling. The thruster and instrumentation designs were proven to be effective via hot fire testing. The thruster diagnostics provided inner wall temperature and static pressure measurements which were compared to the thruster global performance data. For several operating conditions, the performance data exhibited unexpected trends which were correlated with changes in the axial wall temperature distribution. Azimuthal temperature distributions were found to be a function of operating conditions and hardware configuration. The static pressure profiles showed that no severe pressure gradients were present in the rocket. The results indicated that small differences in injector design can result in dramatically different thruster performance and wall temperature behavior, but that these injector effects may be overshadowed by operating at a high fuel film cooling rate.

Despin System for Hydrogen Tank in the Propulsion Systems Laboratory

NASA Image and Video Library

1962-04-21

Mechanic Howard Wine inspects the setup of a spin isolator in Cell 2 of the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Photographer Al Jecko filmed the proceedings. This test was unique in that the chamber’s altitude system was used, but not its inlet air flow. The test was in preparation for an upcoming launch of modified liquid hydrogen propellant tank on a sounding rocket. This Weightlessness Analysis Sounding Probe (WASP) was part of Lewis investigation into methods for controlling partially filled liquid hydrogen fuel tanks during flight. Second-stage rockets, the Centaur in particular, were designed to stop their engines and coast, then restart them when needed. During this coast period, the propellant often shifted inside the tank. This movement could throw the rocket off course or result in the sloshing of fuel away from the fuel pump. Wine was one of only three journeymen mechanics at Lewis when he was hired in January 1954. He spent his first decade in the Propulsion Systems Laboratory and was soon named a section head. Wine went on to serve as Assistant Division Chief and later served as an assistant to the director. Jecko joined the center in 1947 as a photographer and artist. He studied at the Cleveland School or Art and was known for his cartoon drawing. He worked at the center for 26 years.

Ignition of Liquid Fuel Spray and Simulated Solid Rocket Fuel by Photoignition of Carbon Nanotube Utilizing a Camera Flash

DTIC Science & Technology

2011-12-01

10,11 There has been a recent report on the photoignition of graphene oxide for fuel ignition applications.12 In this report, we will describe the...slide Aluminum foil Glass petri dish Xe flash Camera Sample Black spray paint Figure 2- Schematic and photographs of the experimental setup...Gilje, Sergey Dubin, Alireza Badakhshan, Jabari Farrar, Stephen. A. Danczyk, Richard B. Kaner, “Photothermal Deoxygenation of Graphene Oxide for

Vibration, acoustic, and shock design and test criteria for components on the Solid Rocket Boosters (SRB), Lightweight External Tank (LWT), and Space Shuttle Main Engines (SSME)

NASA Technical Reports Server (NTRS)

1984-01-01

The vibration, acoustics, and shock design and test criteria for components and subassemblies on the space shuttle solid rocket booster (SRB), lightweight tank (LWT), and main engines (SSME) are presented. Specifications for transportation, handling, and acceptance testing are also provided.

NTREES Testing and Operations Status

NASA Technical Reports Server (NTRS)

Emrich, Bill

2007-01-01

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today's best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. Currently, the construction of such a simulator has been completed at the Marshall Space Flight Center, and will be used in the future to evaluate a wide variety of fuel element designs and the materials of which they are fabricated. This present work addresses the operational status of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

High altitude chemical release systems for project BIME (Brazilian Ionospheric Modification Experiments) project IMS (Ionospheric Modification Studies) project PIIE (Polar Ionospheric Irregularities Experiment) project polar arcs

NASA Astrophysics Data System (ADS)

Stokes, Charles S.; Murphy, William J.

1987-07-01

Project BIME, a Spread F observation program involved the launching of two Nike-Black Brant rockets each containing a payload of Ammonium Nitrate Fuel Oil (ANFO). The rockets were launched from Barriera Do Inferno Launch Site in Natal, Brazil in August of 1982. Project IMS, an F-layer modification experiment involved three launch vehicles, a Nike-Tomahawk and two Sonda III rockets. The Nike-Tomahawk carried a sulfur hexafluoride (SF6) payload. One of the Sonda III rockets carried a payload that consisted of an SF6 canister and a samarium/strontium thermite canister. The remaining Sonda III carried a trifluorobromo methane (CF3Br) canister and a samarium thermite canister. The rockets were launched from Wallops Island Launch Facility, Virginia in November of 1984. Project PIIE and Polar Arcs, a program to investigate polar ionospheric irregularities, involved a Nike-Black Brant rocket carrying one samarium thermite canister and six barium canisters. An attempted launch failed when launch criteria could not be met. The rocket was launched successfully from Sondrestrom Air Base, Greenland in March 1987.

KSC-2012-4454

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4453

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4448

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on to trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4449

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4450

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4452

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4451

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

Descent Stage of Mars Science Laboratory During Assembly

NASA Technical Reports Server (NTRS)

2008-01-01

This image from early October 2008 shows personnel working on the descent stage of NASA's Mars Science Laboratory inside the Spacecraft Assembly Facility at NASA's Jet Propulsion Laboratory, Pasadena, Calif.

The descent stage will provide rocket-powered deceleration for a phase of the arrival at Mars after the phases using the heat shield and parachute. When it nears the surface, the descent stage will lower the rover on a bridle the rest of the way to the ground. The larger three of the orange spheres in the descent stage are fuel tanks. The smaller two are tanks for pressurant gas used for pushing the fuel to the rocket engines.

JPL, a division of the California Institute of Technology, manages the Mars Science Laboratory Project for the NASA Science Mission Directorate, Washington.

Further analytical study of hybrid rocket combustion

NASA Technical Reports Server (NTRS)

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

1972-01-01

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

KSC-2012-4444

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, preparations are underway to load a twin set of space shuttle solid rocket boosters and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

Flight Testing of Terrain-Relative Navigation and Large-Divert Guidance on a VTVL Rocket

NASA Technical Reports Server (NTRS)

Trawny, Nikolas; Benito, Joel; Tweddle, Brent; Bergh, Charles F.; Khanoyan, Garen; Vaughan, Geoffrey M.; Zheng, Jason X.; Villalpando, Carlos Y.; Cheng, Yang; Scharf, Daniel P.;

-----SPACE TRANSPORTATION

NASA Image and Video Library

1998-10-07

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.

Rocketdyne RBCC Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratckin, G.; Goldman, A.; Ortwerth, P.; Weisberg, S.

1999-01-01

Boeing Rocketdyne is pursuing the development of Rocket Based Combined Cycle (RBCC), propulsion systems as demonstrated by significant contract work in the hypersonic arena (ART, NASP, SCT, system studies) and over 12 years of steady company discretionary investment. The Rocketdyne concept is a fixed geometry integrated rocket, ramjet, scramjet which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals. seal purge gas, and closeout side attachments. Rocketdyne's experimental RBCC engine (Engine A5) was constructed under contract with the NASA Marshall Space Flight Center. Engine A5 models the complete flight engine flowpath consisting of an inlet, isolator, airbreathing combustor and nozzle. High performance rocket thrusters are integrated into the engine to enable both air-augmented rocket (AAR) and pure rocket operation. Engine A5 was tested in CASL's new FAST facility as an air-augmented rocket, a ramjet and a pure rocket. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. Rocket mode performance was above predictions. For the first time. testing also demonstrated transition from AAR operation to ramjet operation. This baseline configuration has also been shown, in previous testing, to perform well in the scramjet mode.

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.

Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

2015-01-01

The Nuclear Thermal Rocket (NTR) represents the next evolutionary step in cryogenic liquid rocket engines. Deriving its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core, the NTR can generate high thrust at a specific impulse of approx. 900 seconds or more - twice that of today's best chemical rockets. In FY'11, as part of the AISP project, NASA proposed a Nuclear Thermal Propulsion (NTP) effort that envisioned two key activities - "Foundational Technology Development" followed by system-level "Technology Demonstrations". Five near-term NTP activities identified for Foundational Technology Development became the basis for the NCPS project started in FY'12 and funded by NASA's AES program. During Phase 1 (FY'12-14), the NCPS project was focused on (1) Recapturing fuel processing techniques and fabricating partial length "heritage" fuel elements for the two candidate fuel forms identified by NASA and the DOE - NERVA graphite "composite" and the uranium dioxide (UO2) in tungsten "cermet". The Phase 1 effort also included: (2) Engine Conceptual Design; (3) Mission Analysis and Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable and Sustainable NTP Development Strategy. During FY'14, a preliminary plan for DDT&E was outlined by GRC, the DOE and industry for NASA HQ that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce development costs, the GTD and FTD tests use a small, low thrust (approx. 7.5 or 16.5 klbf) engine. Both engines use graphite composite fuel and a "common" fuel element design that is scalable to higher thrust (approx. 25 klbf) engines by increasing the number of elements in a larger diameter core that can produce greater thermal power output. To keep the FTD mission cost down, a simple "1-burn" lunar flyby mission was considered along with maximizing the use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10-B2 engine and Delta Cryogenic Second Stage) to further ensure affordability. This paper provides a preliminary NASA, DOE and industry assessment of what is required - the key DDT&E activities, development options, and the associated schedule - to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

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.

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.

Rocketdyne Development of RBCC Engine for Low Cost Access to Space

NASA Technical Reports Server (NTRS)

Ortwerth, P.; Ratekin, G.; Goldman, A.; Emanuel, M.; Ketchum, A.; Horn, M.

1997-01-01

Rocketdyne is pursuing the conceptual design and development of a Rocket Based Combined Cycle (RBCC) engine for booster and SSTO, advanced reusable space transportation ARTT systems under contract with NASA Marshall Space Flight Center. The Rocketdyne concept is fixed geometry integrated Rocket, Ram Scramjet which is Hydrogen fueled and uses Hydrogen regenerative cooling. Vision vehicle integration studies have determined that scramjet operation to Mach 12 has high payoff for low cost reusable space transportation. Rocketdyne is internally developing versions of the concept for other applications in high speed aircraft and missiles with Hydrocarbon fuel systems. Subscale engine ground testing is underway for all modes of operation from takeoff to Mach 8. High altitude Rocket only mode tests will be completed as part of the ground test program to validate high expansion ratio performance. A unique feature of the ground test series is the inclusion of dynamic trajectory simulation with real time Mach number, altitude, engine throttling, and RBCC mode changes in a specially modified freejet test facility at GASL. Preliminary cold flow Air Augmented Rocket mode test results and Short Combustor tests have met program goals and have been used to integrate all modes of operation in a single combustor design with a fixed geometry inlet for design confirmation tests. A water cooled subscale engine is being fabricated and installed for test beginning the last quarter of 1997.

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

NASA Technical Reports Server (NTRS)

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

2017-01-01

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

Aircraft Research and Technology for Antimisting Kerosene Conference, February 18-19, 1981.

DTIC Science & Technology

1981-06-01

carrier turbine aircraft fatal accidents from 1964 through 1976. Since antimisting fuel is intended to inhibit ignition and flame propagation when fuel is...been shown to be possible and rapid, although costly and complex. One item that should be added at this point is in the event turbine engine power...port side with 0.28 percent FM-9 fuel was ignited by the rocket motors. When the turbine engine separated from the wing, localized fire remained with

Project Summary: Increased Fuel Affordability Through Deployable Refining Technology (Briefing Charts)

DTIC Science & Technology

2016-05-01

distribution is unlimited. PA Clearance #16274 2 Overall Process DISTRIBUTION A. Approved for public release; distribution unlimited. 3 Sulfur in Fuels...Mercaptans (Thiols) The presence of sulfur in fuels leads to many detrimental effects: • Coking on rocket engine injector plates • Deposit formation in... Sulfur Compounds by GC-SCD ( Sulfur Speciation) Concentration (ppm) C2 Thiophenes 0.3 C3-C4 Thiophenes 1.4 C5 Thiophenes 3.7 C6 Thiophenes 3.5 C7

Spreading the Word on Safety

NASA Technical Reports Server (NTRS)

2004-01-01

Beginning with the Apollo Program in the early 1960s, the NASA White Sands Test Facility (WSTF) has supported every U.S. human exploration space flight program to date. Located in Las Cruces, New Mexico, WSTF is part of Johnson Space Center. The facility's primary mission is to provide the expertise and infrastructure to test and evaluate spacecraft materials, components, and rocket propulsion systems to enable the safe human exploration and utilization of space. WSTF stores, tests, and disposes of Space Shuttle and International Space Station propellants. Since aerospace fluids can have harmful reactions with the construction materials of the systems containing them, a major component of WSTF's work is the study of propellants and hazardous materials. WSTF has a wide variety of resources to draw upon in assessing the fire, explosion, compatibility, and safety hazards of these fluids, which include hydrogen, oxygen, hydrazine fuels, and nitrogen tetroxide. In addition to developing new test methods, WSTF has created technical manuals and training courses for the safe use of aerospace fluids.

Hydrogen/Air Fuel Nozzle Emissions Experiments

NASA Technical Reports Server (NTRS)

Smith, Timothy D.

2001-01-01

The use of hydrogen combustion for aircraft gas turbine engines provides significant opportunities to reduce harmful exhaust emissions. Hydrogen has many advantages (no CO2 production, high reaction rates, high heating value, and future availability), along with some disadvantages (high current cost of production and storage, high volume per BTU, and an unknown safety profile when in wide use). One of the primary reasons for switching to hydrogen is the elimination of CO2 emissions. Also, with hydrogen, design challenges such as fuel coking in the fuel nozzle and particulate emissions are no longer an issue. However, because it takes place at high temperatures, hydrogen-air combustion can still produce significant levels of NOx emissions. Much of the current research into conventional hydrocarbon-fueled aircraft gas turbine combustors is focused on NOx reduction methods. The Zero CO2 Emission Technology (ZCET) hydrogen combustion project will focus on meeting the Office of Aerospace Technology goal 2 within pillar one for Global Civil Aviation reducing the emissions of future aircraft by a factor of 3 within 10 years and by a factor of 5 within 25 years. Recent advances in hydrocarbon-based gas turbine combustion components have expanded the horizons for fuel nozzle development. Both new fluid designs and manufacturing technologies have led to the development of fuel nozzles that significantly reduce aircraft emissions. The goal of the ZCET program is to mesh the current technology of Lean Direct Injection and rocket injectors to provide quick mixing, low emissions, and high-performance fuel nozzle designs. An experimental program is planned to investigate the fuel nozzle concepts in a flametube test rig. Currently, a hydrogen system is being installed in cell 23 at NASA Glenn Research Center's Research Combustion Laboratory. Testing will be conducted on a variety of fuel nozzle concepts up to combustion pressures of 350 psia and inlet air temperatures of 1200 F. Computational fluid dynamics calculations, with the Glenn developed National Combustor Code, are being performed to optimize the fuel nozzle designs.

Black carbon cookstove emissions: A field assessment of 19 stove/fuel combinations

NASA Astrophysics Data System (ADS)

Garland, Charity; Delapena, Samantha; Prasad, Rajendra; L'Orange, Christian; Alexander, Donee; Johnson, Michael

2017-11-01

Black carbon (BC) emissions from household cookstoves consuming solid fuel produce approximately 25 percent of total anthropogenic BC emissions. The short atmospheric lifetime of BC means that reducing BC emissions would result in a faster climate response than mitigating CO2 and other long-lived greenhouse gases. This study presents the results of optical BC measurements of two new cookstove emissions field assessments and 17 archived cookstove datasets. BC was determined from attenuation of 880 nm light, which is strongly absorbed by BC, and linearly related between 1 and 125 attenuation units. A relationship was experimentally determined correlating BC mass deposition on quartz filters determined via thermal optical analysis (TOA) and on PTFE and quartz filters using transmissometry, yielding an attenuation cross-section (σATN) for both filter media types. σATN relates TOA measurements to optical measurements on PTFE and quartz (σATN(PTFE) = 13.7 cm-2 μg, R2 = 0.87, σATN(Quartz) = 15.6 cm-2 μg, R2 = 0.87). These filter-specific σATN, optical measurements of archived filters were used to determine BC emission factors and the fraction of particulate matter (PM) in the form of black carbon (BC/PM). The 19 stoves measured fell into five stove classes; simple wood, rocket, advanced biomass, simple charcoal, and advanced charcoal. Advanced biomass stoves include forced- and natural-draft gasifiers which use wood or biomass pellets as fuel. Of these classes, the simple wood and rocket stoves demonstrated the highest median BC emission factors, ranging from 0.051 to 0.14 g MJ-1. The lowest BC emission factors were seen in charcoal stoves, which corresponds to the generally low PM emission factors observed during charcoal combustion, ranging from 0.0084 to 0.014 g MJ-1. The advanced biomass stoves generally showed an improvement in BC emissions factors compared to simple wood and rocket stoves, ranging from 0.0031 to 0.071 g MJ-1. BC/PM ratios were highest for the advanced and rocket stoves. Potential relative climate impacts were estimated by converting aerosol emissions to CO2-equivalent, and suggest that some advanced stove/fuel combinations could provide substantial climate benefits.

Saturn Apollo Program

NASA Image and Video Library

1962-11-16

The Saturn I (SA-3) flight lifted off from Kennedy Space Center launch Complex 34, November 16, 1962. The third launch of Saturn launch vehicles, developed at the Marshall Space Flight Center (MSFC) under the direction of Dr. Wernher von Braun, incorporated a Saturn I, Block I engine. The typical height of a Block I vehicle was approximately 163 feet. and had only one live stage. It consisted of eight tanks, each 70 inches in diameter, clustered around a central tank, 105 inches in diameter. Four of the external tanks were fuel tanks for the RP-1 (kerosene) fuel. The other four, spaced alternately with the fuel tanks, were liquid oxygen tanks as was the large center tank. All fuel tanks and liquid oxygen tanks drained at the same rates respectively. The thrust for the stage came from eight H-1 engines, each producing a thrust of 165,000 pounds, for a total thrust of over 1,300,000 pounds. The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis and canted outward slightly, while the remaining four engines were located outboard in a larger square pattern offset 40 degrees from the inner pattern. Unlike the inner engines, each outer engine was gimbaled. That is, each could be swung through an arc. They were gimbaled as a means of steering the rocket, by letting the instrumentation of the rocket correct any deviations of its powered trajectory. The block I required engine gimabling as the only method of guiding and stabilizing the rocket through the lower atmosphere. The upper stages of the Block I rocket reflected the three-stage configuration of the Saturn I vehicle. During the SA-3 flight, the upper stage ejected 113,560 liters (30,000 gallons) of ballast water in the upper atmosphere for "Project Highwater" physics experiment. The water was released at an altitude of 65 miles, where within only 5 seconds, it expanded into a massive ice cloud 4.6 miles in diameter. Release of this vast quantity of water in a near-space environment marked the first purely scientific large-scale experiment.

Space propulsion systems. Present performance limits and application and development trends

NASA Technical Reports Server (NTRS)

Buehler, R. D.; Lo, R. E.

1981-01-01

Typical spaceflight programs and their propulsion requirements as a comparison for possible propulsion systems are summarized. Chemical propulsion systems, solar, nuclear, or even laser propelled rockets with electrical or direct thermal fuel acceleration, nonrockets with air breathing devices and solar cells are considered. The chemical launch vehicles have similar technical characteristics and transportation costs. A possible improvement of payload by using air breathing lower stages is discussed. The electrical energy supply installations which give performance limits of electrical propulsion and the electrostatic ion propulsion systems are described. The development possibilities of thermal, magnetic, and electrostatic rocket engines and the state of development of the nuclear thermal rocket and propulsion concepts are addressed.

GOES-S Countdown to T-Zero, Episode 3: Rocket Science

NASA Image and Video Library

2018-02-27

The United Launch Alliance Atlas V rocket reaches another major milestone on the road to T-Zero, as NOAA's GOES-S spacecraft prepares for launch. Stacking the rocket begins with the booster - the largest component - and continues with the addition of four solid rocket motors and the Centaur upper stage. GOES-S, the next in a series of advanced weather satellites, is slated to launch aboard the Atlas V from Cape Canaveral Air Force Station in Florida.

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.

Air Force Research Laboratory (AFRL) research highlights, September--October 1998

DOE Office of Scientific and Technical Information (OSTI.GOV)

NONE

New AFOSR-sponsored research shows that exhausts from solid-fueled rocket motors have very limited impact on stratospheric ozone. The research provides the Air Force with hard data to support continued access to space using the existing fleet of rockets and rocket technology. This basic research data allows the Air Force to maintain a strongly proactive environmental stance, and to meet federal guidelines regarding environmental impacts. Long-standing conjecture within the international rocket community suggests that chlorine compounds and alumina particulates produced in solid rocket motor (SRM) exhausts could create localized, temporary ozone toss in rocket plumes following launches. The extent of amore » local depletion of ozone and its environmental impact depends on details of the composition and chemistry in these plumes. Yet direct measurements of plume composition and plume chemistry in the stratosphere had never been made. Uncertainty about these details left the Air Force and commercial space launch capability potentially vulnerable to questions about the environmental impact of rocket launches. In 1995, APOSR and the Space and Missiles Systems Center Launch Programs Office (SMC/CL) jointly began the Rocket Impacts on Stratospheric Ozone (RISO) program to make the first-ever detailed measurements of rocket exhaust plumes. These measurements were aimed at understanding how the exhaust from large rocket motors effect the Earth`s stratospheric ozone layer. The studies determined: the size distribution of alumina particles in these exhausts, the amount of reactive chlorine in SRM exhaust, and the size and duration of localized ozone toss in the rocket plumes.« less

Low-Cost, High-Performance Cryocoolers for In-Situ Propellant Production

NASA Technical Reports Server (NTRS)

Martin, J. L.; Corey, J. A.; Peters, T. A.

1999-01-01

A key feature of many In-Situ Resource Utilization (ISRU) schemes is the production of rocket fuel and oxidizer from the Martian atmosphere. Many of the fuels under consideration will require cryogenic cooling for efficient long-term storage. Although significant research has been focused on the techniques for producing the fuels from Martian resources, little effort has been expended on the development of cryocoolers to efficiently liquefy these fuels. This paper describes the design of a pulse tube liquefier optimized for liquefying oxygen produced by an In-Situ Propellant Production (ISPP) plant on Mars.

Low-Cost High-Performance Cryocoolers for In-Situ Propellant Production

NASA Technical Reports Server (NTRS)

Martin, J. L.; Corey, J. A.; Peters, T. A.

1999-01-01

A key feature of many In-Situ Resource Utilization (ISRU) schemes is the production of rocket fuel and oxidizer from the Martian atmosphere. Many of the fuels under consideration will require cryogenic cooling for efficient long-term storage. Although significant research has been focused on the techniques for producing the fuels from Martian resources, little effort has been expended on the development of cryocoolers to efficiently liquefy these fuels. This paper describes the design of a pulse tube liquefier optimized for liquefying oxygen produced by an In-Situ Propellant Production (ISPP) plant on Mars.

Bimodal Nuclear Thermal Rocket Sizing and Trade Matrix for Lunar, Near Earth Asteroid and Mars Missions

NASA Astrophysics Data System (ADS)

McCurdy, David R.; Krivanek, Thomas M.; Roche, Joseph M.; Zinolabedini, Reza

2006-01-01

The concept of a human rated transport vehicle for various near earth missions is evaluated using a liquid hydrogen fueled Bimodal Nuclear Thermal Propulsion (BNTP) approach. In an effort to determine the preliminary sizing and optimal propulsion system configuration, as well as the key operating design points, an initial investigation into the main system level parameters was conducted. This assessment considered not only the performance variables but also the more subjective reliability, operability, and maintainability attributes. The SIZER preliminary sizing tool was used to facilitate rapid modeling of the trade studies, which included tank materials, propulsive versus an aero-capture trajectory, use of artificial gravity, reactor chamber operating pressure and temperature, fuel element scaling, engine thrust rating, engine thrust augmentation by adding oxygen to the flow in the nozzle for supersonic combustion, and the baseline turbopump configuration to address mission redundancy and safety requirements. A high level system perspective was maintained to avoid focusing solely on individual component optimization at the expense of system level performance, operability, and development cost.

Friction Plug Weld Repair for the Space Shuttle External Tank

NASA Technical Reports Server (NTRS)

Hartley, Paula J.; McCool, A. (Technical Monitor)

2000-01-01

Lockheed Martin Space Systems, Michoud Operations in New Orleans, LA is the manufacturer of the External Fuel Tanks (ET) for the Space Transportation System (STS). The ET contains and delivers the propellants used by the Orbiters three main engines. Additionally, it also serves as the structural backbone for the Orbiter and the two Solid Rocket Boosters (SRB), which combined, constitute the STS. In 1994, NASA established that in order to launch the International Space Station, the performance of the STS must be improved. One option was to reduce the weight of the ET, which would enable sufficient increase in performance. With the development of the Weldalite(R) series of Al-Cu-Li alloys in the late 1980's, Lockheed Martin was postured to replace the current A12219 fuel tanks with the high strength, light weight A12195 alloy. With the use of A12195 and some component redesign, the weight of the Super Lightweight (SLWT) ET was reduced by approximately 7,000 pounds, which added as much capability to the Space Shuttle. Since June 1998, seven STS missions have been successful with the use of the SLWT ET's.

Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Corban, Robert R.; Mcguire, Melissa L.; Beke, Erik G.

1995-01-01

The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLLV) and smaller 120 t HLLV option. Attractive performance characteristics and high-leverage technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for commonality of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed.

Infrasound and seismic analysis of the SpaceX Falcon9 explosion sequence of 1-September-2016

NASA Astrophysics Data System (ADS)

Thompson, G.; McNutt, S. R.; Brown, R. G.; Braunmiller, J.; Mehta, C.

2017-12-01

During a static launch test on 1-Sep-2016 at Kennedy Space Center, a SpaceX Falcon 9 rocket exploded causing loss of the rocket and the payload, and extensively damaging the launch complex. The sequence was captured by a 3-element infrasound array and a broadband 3-component seismometer at the Astronaut Beach House, just 0.87 miles (1.4 km) from the launch pad. Manual picking identified 153 impulsive airwave signals over a 26-minute interval and these were compared to video recordings of the sequence. The explosion onset consisted of a moderate signal on both seismic and infrasound (52 Pa) instruments. This corresponds to the rupture of the second-stage fuel tank. We found no signals before this, so we do not believe that there was an external cause. The primary fuel tank ruptured 4 seconds later and was the strongest event by far, producing an infrasound signal that exceeded 1400 Pa ( 2000 Pa in reduced pressure). The seismic signal consists mainly of air-coupled Rayleigh waves with frequencies of 5-23 Hz. The infrasound events occurred in four clusters. The first cluster included the onset and main events and 46 smaller events. This was followed by several minutes without infrasound signals during which a 3.5 minute continuous seismic vibration occurred. Cluster 2 consisted of 4 events ranging from 117-256 Pa. Cluster 3 comprised 96 events of 7-78 Pa. Cluster 4 consisted of 5 events with overpressures of 23-63 Pa. Gaps of several minutes without infrasound and seismic signals occurred between clusters 2 and 3, and 3 and 4. In terms of energy, the main event dominated; in terms of numbers, cluster 3 had the most infrasound events. The seismic and infrasound data are complementary to video recordings of the explosion, and provide additional characterization that may be useful to interpret the sequence of events. Because of the proximity of our array to this rocket explosion, our dataset may be unique.

Unique thermocouple to measure the temperatures of squibs, igniters, propellants, and rocket nozzles

NASA Astrophysics Data System (ADS)

Nanigian, Jacob; Nanigian, Dan

2006-05-01

The temperatures produced by the various components in the propulsion system of rockets and missiles determine the performance of the rocket. Since these temperatures occur very rapidly and under extreme conditions, standard thermocouples fail before any meaningful temperatures are measured. This paper describes the features of a special family of high performance thermocouples, which can measure these transient temperatures with millisecond response times and under the most severe conditions of erosion. Examples of igniter, propellant and rocket nozzle temperatures are included in this paper. Also included is heat flux measurements made by these sensors in rocket applications.

KSC-2009-5811

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – At Port Canaveral, a tugboat relieves the Liberty Star for the move of the Pegasus barge, carrying external tank 134, through the Banana River toward the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by the solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5812

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – At Port Canaveral, a tugboat begins to push the Pegasus barge, carrying external tank 134, through the Banana River toward the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

KSC-2009-5813

NASA Image and Video Library

2009-10-24

CAPE CANAVERAL, Fla. – At Port Canaveral, a tugboat begins to move the Pegasus barge, carrying external tank 134, through the Banana River toward the turn basin near the Vehicle Assembly Building, or VAB, at NASA's Kennedy Space Center in Florida. Pegasus arrived in Florida after an ocean voyage towed by a solid rocket booster retrieval ship from NASA's Michoud Assembly Facility near New Orleans. After Pegasus docks in the turn basin, the fuel tank will be offloaded and transported into the VAB. ET-134 will be used to launch space shuttle Endeavour on the STS-130 mission to the International Space Station. Launch is targeted for Feb. 4, 2010. For information on the components of the space shuttle and the STS-130 mission, visit http://www.nasa.gov/shuttle. Photo credit: NASA/Jack Pfaller

STS-2 - SOFTWARE INTEGRATION TESTS (SIT) - KSC

NASA Image and Video Library

1981-09-01

S81-36331 (24 Aug. 1981) --- Astronauts Joe H. Engle, left, and Richard H. Truly pause before participating in the integrated test of the assembled space shuttle components scheduled for launch no earlier than Sept. 30, 1981. Moments later, Engle, STS-2 crew commander, and Truly, pilot, entered the cabin of the orbiter Columbia for a mission simulation. The shuttle integrated tests (SIT) are designed to check out every connection and signal path in the STS-2 vehicle composed of the orbiter, two solid rocket boosters (SRB) and an external fuel tank (ET) for Columbia?s main engines. Completion of the tests will clear the way for preparations for rollout to Pad A at Launch Complex 39, scheduled for the latter part of August or early September. Photo credit: NASA

The E-3 Test Facility at Stennis Space Center: Research and Development Testing for Cryogenic and Storable Propellant Combustion Systems

NASA Technical Reports Server (NTRS)

Pazos, John T.; Chandler, Craig A.; Raines, Nickey G.

2009-01-01

This paper will provide the reader a broad overview of the current upgraded capabilities of NASA's John C. Stennis Space Center E-3 Test Facility to perform testing for rocket engine combustion systems and components using liquid and gaseous oxygen, gaseous and liquid methane, gaseous hydrogen, hydrocarbon based fuels, hydrogen peroxide, high pressure water and various inert fluids. Details of propellant system capabilities will be highlighted as well as their application to recent test programs and accomplishments. Data acquisition and control, test monitoring, systems engineering and test processes will be discussed as part of the total capability of E-3 to provide affordable alternatives for subscale to full scale testing for many different requirements in the propulsion community.

14 CFR 437.33 - Landing and impact locations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... each location for nominal landing or any contingency abort landing of the reusable suborbital rocket, and each location for any nominal or contingency impact or landing of a component of that rocket...

14 CFR 437.33 - Landing and impact locations.

Code of Federal Regulations, 2012 CFR

2012-01-01

... each location for nominal landing or any contingency abort landing of the reusable suborbital rocket, and each location for any nominal or contingency impact or landing of a component of that rocket...

14 CFR 437.33 - Landing and impact locations.

Code of Federal Regulations, 2013 CFR

2013-01-01

... each location for nominal landing or any contingency abort landing of the reusable suborbital rocket, and each location for any nominal or contingency impact or landing of a component of that rocket...

14 CFR 437.33 - Landing and impact locations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... each location for nominal landing or any contingency abort landing of the reusable suborbital rocket, and each location for any nominal or contingency impact or landing of a component of that rocket...

14 CFR 437.33 - Landing and impact locations.

Code of Federal Regulations, 2014 CFR

2014-01-01

... each location for nominal landing or any contingency abort landing of the reusable suborbital rocket, and each location for any nominal or contingency impact or landing of a component of that rocket...

United States Air Force Graduate Student Summer Support Program (1987). Program Technical Report. Volume 1.

DTIC Science & Technology

1987-12-01

developed for a large percentage of the participants in the Summer Faculty Research Program in 1979-1983 period through an AFOSR Minigrant Program . On 1...Analysis of a Bimodal Nuclear Rocket Core by Dav,, C. Carpenter ABSTRACT The framework for a general purpose finite element analysis code was developed ...to study the 2-D temperature distribution in a hot-channel S hexagonal fuel element in the core of a bimodal nuclear’ rocket. Prelim- inary thermal

Fuel Chemistry and Combustion Distribution Effects on Rocket Engine Combustion Stability

DTIC Science & Technology

2012-01-25

by Crowe et al. (1963). The small solid rocket motors are fired into the collection tank with the nozzle [Crowe et al. (1963)] and without nozzle...explosions at the end of the droplet lifetime. Upon ignition , a neat droplet of JP-8 will burn orange, and the droplet will regress until all of the...pixel location were estimated by applying a time shift and amplitude scaling factor to the pressure measurements made at the aft end of chamber

Toward Active Control of Noise from Hot Supersonic Jets

DTIC Science & Technology

2014-04-21

regions of the jet. A retro -reflective shadowgraph setup was used to record the images. The near-nozzle region exhibits a large number of shock-like...jet exit plane; nearly identical observations have been made in the rocket noise community [15, 29| . The only discrepancies in figure 9b are with the...noise surveys of solid-fuel rocket engines for a range of nozzle exit pressures," NASA TN D-21, August, 1959. [16] Potter, R.C. and Jones, J.H., "An

Technology Innovations from NASA's Next Generation Launch Technology Program

NASA Technical Reports Server (NTRS)

Cook, Stephen A.; Morris, Charles E. K., Jr.; Tyson, Richard W.

2004-01-01

NASA's Next Generation Launch Technology Program has been on the cutting edge of technology, improving the safety, affordability, and reliability of future space-launch-transportation systems. The array of projects focused on propulsion, airframe, and other vehicle systems. Achievements range from building miniature fuel/oxygen sensors to hot-firings of major rocket-engine systems as well as extreme thermo-mechanical testing of large-scale structures. Results to date have significantly advanced technology readiness for future space-launch systems using either airbreathing or rocket propulsion.

Russia and NATO Missile Defense: The European Phased Adaptive Approach Experience, 2009-2017

DTIC Science & Technology

2018-03-01

the Manhattan Project, the U.S. effort to develop the atomic bomb . Donald L. Hafner observes, “In 1941, two years after crucial scientific work by...Szilard and Fermi suggested an atomic bomb might be feasible, the Manhattan Project began its task with promise from the project advocates that a...German V-2 rocket. The V-1 flying bomb preceded the V-2 rocket and was only somewhat unreliable but was able to save on both fuel and air crews.35

An inverse dynamics approach to trajectory optimization for an aerospace plane

NASA Technical Reports Server (NTRS)

Lu, Ping

1992-01-01

An inverse dynamics approach for trajectory optimization is proposed. This technique can be useful in many difficult trajectory optimization and control problems. The application of the approach is exemplified by ascent trajectory optimization for an aerospace plane. Both minimum-fuel and minimax types of performance indices are considered. When rocket augmentation is available for ascent, it is shown that accurate orbital insertion can be achieved through the inverse control of the rocket in the presence of disturbances.

History of the Redstone Missile System

DTIC Science & Technology

1965-10-15

knowledge in areas such as propulsion systems, rocket fuels, aerodynamics, guidance equipment, and’testing equipment. It compiled basic statistics on...i n g of t he 46th and 4 3 Tech Rept, ABMA, 30 Jun 57, sub: Ordnance Guided M i s s i l e and Rocket Programs, Redstone, Vol. IVg Supp. 2, p. 67...embarked in June for Europe. The main body boarded ship on 18 June 1958 for Saint-Nazaire, France , and moved in convoy across France and 48~ept

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.

NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

NASA Technical Reports Server (NTRS)

Arrighi, Robert S.

2016-01-01

"There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best Oklahoma oil field tradition." Besides the Rocket Systems Area, Plum Brook Station also included a nuclear test reactor, a large vacuum tank, a hypersonic wind tunnel, and a full-scale upper-stage rocket stand. The Rocket Systems Area operated from 1961 until NASA shut down all of Plum Brook in 1974. The center reopened Plum Brook in the late 1980s and continues to use several test facilities. The Rocket Systems Area, however, was not restored. Today Plum Brook resembles a nature preserve more than an oil refinery. Lush fields and forests separate the large test facilities. Until recently, the abandoned Rocket Systems Area structures and equipment were visible amongst the greenery. These space-age ruins, particularly the three towers, stood as silent sentinels over the sparsely populated reservation. Few knew the story of these mysterious facilities when NASA removed them in the late 2000s.

SOUTHEAST AND NORTHEAST SIDES. Looking west Edwards Air Force ...

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

SOUTHEAST AND NORTHEAST SIDES. Looking west - Edwards Air Force Base, Air Force Rocket Propulsion Laboratory, Fuel & Water Tank, Test Area 1-115, northwest end of Saturn Boulevard, Boron, Kern County, CA

Ultrasonic Inspection Of Thick Sections

NASA Technical Reports Server (NTRS)

Friant, C. L.; Djordjevic, B. B.; O'Keefe, C. V.; Ferrell, W.; Klutz, T.

1993-01-01

Ultrasonics used to inspect large, relatively thick vessels for hidden defects. Report based on experiments in through-the-thickness transmission of ultrasonic waves in both steel and filament-wound composite cases of solid-fuel rocket motors.

KSC-2012-4442

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a twin set of space shuttle solid rocket boosters and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4438

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a twin set of space shuttle solid rocket boosters and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

Development of 90 kgf Class CAMUI Hybrid Rocket for a CanSat Experiment

NASA Astrophysics Data System (ADS)

Nagata, Harunori; Uematsu, Tsutomu; Ito, Mitsunori; Kakikura, Akihito; Kaneko, Yudai; Mori, Kazuhiro; Murai, Norikazu; Sato, Tatsuhiro; Mitsuhashi, Ryuichi; Totani, Tsuyoshi

A newly designed CAMUI hybrid rocket motor of 900 N (90 kgf) thrust class, CAMUI-90, was developed. It uses a combination of polyethylene and liquid oxygen as propellants. CAMUI hybrid rocket is an explosive-flee small rocket motor to realize a small launch system with low cost and flexibility. The motor produces a thrust of 900 N for four seconds, keeping the optimal characteristic exhaust velocity of the fuel-oxidizer combination (exceeding 1800 m/s). A main application of the CAMUI-90 motor is for a CanSat experiment. A launch vehicle employing CAMUI-90 motor, 120 mm in diameter and 3.05 m in length, accelerates a payload of 500 g to 140 m/s in four seconds and reaches to an altitude of about 1 km. The first launch of this vehicle was on December 2006.

Predicting the velocity and azimuth of fragments generated by the range destruction or random failure of rocket casings and tankage

NASA Astrophysics Data System (ADS)

Eck, M.; Mukunda, M.

The proliferation of space vehicle launch sites and the projected utilization of these facilities portends an increase in the number of on-pad, ascent, and on-orbit solid-rocket motor (SRM) casings and liquid-rocket tanks which will randomly fail or will fail from range destruct actions. Beyond the obvious safety implications, these failures may have serious resource implications for mission system and facility planners. SRM-casing failures and liquid-rocket tankage failures result in the generation of large, high velocity fragments which may be serious threats to the safety of launch support personnel if proper bunkers and exclusion areas are not provided. In addition, these fragments may be indirect threats to the general public's safety if they encounter hazardous spacecraft payloads which have not been designed to withstand shrapnel of this caliber. They may also become threats to other spacecraft if, by failing on-orbit, they add to the ever increasing space-junk collision cross-section. Most prior attempts to assess the velocity of fragments from failed SRM casings have simply assigned the available chamber impulse to available casing and fuel mass and solved the resulting momentum balance for velocity. This method may predict a fragment velocity which is high or low by a factor of two depending on the ratio of fuel to casing mass extant at the time of failure. Recognizing the limitations of existing methods, the authors devised an analytical approach which properly partitions the available impulse to each major system-mass component. This approach uses the Physics International developed PISCES code to couple the forces generated by an Eulerian modeled gas flow field to a Lagrangian modeled fuel and casing system. The details of a predictive analytical modeling process as well as the development of normalized relations for momentum partition as a function of SRM burn time and initial geometry are discussed in this paper. Methods for applying similar modeling techniques to liquid-tankage-over-pressure failures are also discussed. These methods have been calibrated against observed SRM ascent failures and on-orbit tankage failures. Casing-quadrant sized fragments with velocities exceeding 100 m/s resulted from Titan 34D-SRM range destruct actions at 10 s mission elapsed time (MET). Casing-quadrant sized fragments with velocities of approx. 200 m/s resulted from STS-SRM range destruct actions at 110 s MET. Similar sized fragments for Ariane third stage and Delta second stage tankage were predicted to have maximum velocities of 260 and 480 m/s respectively. Good agreement was found between the predictions and observations for five specific events and it was concluded that the methods developed have good potential for use in predicting the fragmentation process of a number of generically similar casing and tankage systems.

Design, construction, test and field support of a containerless payload package for rocket flight. [electromagnetic heating and confinement

NASA Technical Reports Server (NTRS)

1977-01-01

The performance of a device for electromagnetically heating and positioning containerless melts during space processing was evaluated during a 360 second 0-g suborbital sounding rocket flight. Components of the electromagnetic containerless processing package (ECPP), its operation, and interface with the rocket are described along with flight and qualification tests results.

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

NASA Astrophysics Data System (ADS)

Waung, Timothy S.

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

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.

Control Room at the NACA’s Rocket Engine Test Facility

NASA Image and Video Library

1957-05-21

Test engineers monitor an engine firing from the control room of the Rocket Engine Test Facility at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Rocket Engine Test Facility, built in the early 1950s, had a rocket stand designed to evaluate high-energy propellants and rocket engine designs. The facility was used to study numerous different types of rocket engines including the Pratt and Whitney RL-10 engine for the Centaur rocket and Rocketdyne’s F-1 and J-2 engines for the Saturn rockets. The Rocket Engine Test Facility was built in a ravine at the far end of the laboratory because of its use of the dangerous propellants such as liquid hydrogen and liquid fluorine. The control room was located in a building 1,600 feet north of the test stand to protect the engineers running the tests. The main control and instrument consoles were centrally located in the control room and surrounded by boards controlling and monitoring the major valves, pumps, motors, and actuators. A camera system at the test stand allowed the operators to view the tests, but the researchers were reliant on data recording equipment, sensors, and other devices to provide test data. The facility’s control room was upgraded several times over the years. Programmable logic controllers replaced the electro-mechanical control devices. The new controllers were programed to operate the valves and actuators controlling the fuel, oxidant, and ignition sequence according to a predetermined time schedule.

14 CFR 437.61 - Landing and impact locations.

Code of Federal Regulations, 2014 CFR

2014-01-01

... locations. For a nominal or any contingency abort landing of a reusable suborbital rocket, or for any nominal or contingency impact or landing of a component of that rocket, a permittee must use a location...

14 CFR 437.61 - Landing and impact locations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... locations. For a nominal or any contingency abort landing of a reusable suborbital rocket, or for any nominal or contingency impact or landing of a component of that rocket, a permittee must use a location...

14 CFR 437.61 - Landing and impact locations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... locations. For a nominal or any contingency abort landing of a reusable suborbital rocket, or for any nominal or contingency impact or landing of a component of that rocket, a permittee must use a location...

14 CFR 437.61 - Landing and impact locations.

Code of Federal Regulations, 2013 CFR

2013-01-01

... locations. For a nominal or any contingency abort landing of a reusable suborbital rocket, or for any nominal or contingency impact or landing of a component of that rocket, a permittee must use a location...

14 CFR 437.61 - Landing and impact locations.

Code of Federal Regulations, 2012 CFR

2012-01-01

... locations. For a nominal or any contingency abort landing of a reusable suborbital rocket, or for any nominal or contingency impact or landing of a component of that rocket, a permittee must use a location...

Advanced materials for radiation-cooled rockets

NASA Technical Reports Server (NTRS)

Reed, Brian; Biaglow, James; Schneider, Steven

1993-01-01

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