Search NASA⌕ Search

SEARCH · Search NASA

Results for “ROCKET THRUST”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Dynamic Analysis of a Building Under Rocket Engine Plume Acoustic Load

Studies have been performed to develop finite-element modeling and simulation techniques to predict the dynamic structural response of Building 4010 to the acoustic load from the plume of high-thrust rocket motors. The building is the Test Control Center and general office space for the E-complex at Stennis Space Center. It is a large single span; light-structured building located approximately 1,000 feet from the E-1 test stand. A three-dimensional shell/beam combined model of the building was built using Pro/Engineer platform and imported into Pro/Mechanica for analysis. An Equivalent Shell technique was developed to simplify the highly complex building structure so that the calculation is more efficient and accurate. A deterministic approach was used for the dynamic analysis. A pre-stressed modal analysis was performed to simulate the weight stiffening of the structure, through which about 200 modes ranging from 0 to 35 Hz were identified. In an initial dynamic frequency analysis, the maximum response over the model was found. Then the complete 3-D distributions of the displacement, as well as the stresses, were calculated through a final frequency analysis. The results were compared to a strain gage and accelerometer recordings from rocket engine tests and showed reasonable agreement.

Qian, Z.↗
NASA NTRS1948

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

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

Bellman, Donald R↗
NASA NTRS1991

The development of a fiber optic Raman temperature measurement system for rocket flows

A fiber-optic Raman diagnostic system for H2/O2 rocket flows is currently under development. This system is designed for measurements of temperature and major species concentration in the combustion chamber and part of the nozzle of a 100 Newton thrust rocket currently undergoing tests. This paper describes a measurement system based on the spontaneous Raman scattering phenomenon. An analysis of the principles behind the technique is given. Software is developed to measure temperature and major species concentration by comparing theoretical Raman scattering spectra with experimentally obtained spectra. Equipment selection and experimental approach are summarized. This experimental effort is part of a program, which is in progress, to evaluate Navier-Stokes based analyses for this class of rockets.

De Groot, Wim A.↗
NASA NTRS1992

The Development of a Fiber Optic Raman Temperature Measurement System for Rocket Flows

A fiberoptic Raman diagnostic system for H2/O2 rocket flows is currently under development. This system is designed for measurement of temperature and major species concentration in the combustion chamber and part of the nozzle of a 100 Newton thrust rocket currently undergoing testing. This paper describes a measurement system based on the spontaneous Raman scattering phenomenon. An analysis of the principles behind the technique is given. Software is developed to measure temperature and major species concentrations by comparing theoretical Raman scattering spectra with experimentally obtained spectra. Equipment selection and experimental approach are summarized. This experimental program is part of a program, which is in progress, to evaluate Navier-Stokes based analyses for this class of rocket.

Degroot, Wim A.↗
NASA NTRS1990

Nuclear thermal rockets using indigenous extraterrestrial propellants

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

Zubrin, Robert M.↗
NASA NTRS1973

Reliability and effective thermal conductivity of three metallic-ceramic composite insulating coatings on cooled hydrogen-oxygen rockets

An experimental investigation of the structural integrity and effective thermal conductivity of three metallic-ceramic composite coatings was conducted. These coatings were plasma sprayed onto the combustion side of water-cooled, 12.7-centimeter throat diameter, hydrogen-oxygen rocket thrust chambers operating at 2.07 to 4.14 meganewtons per square meter chamber pressure. The metallic-ceramic composites functioned for six to 17 cycles and for as long as 213 seconds of rocket operations and could have probably provided their insulating properties for many additional cycles. The effective thermal conductivity of all the coatings was in the range of 0.7472 to 4.483 w/(m)(K), which makes the coatings a very effective thermal barrier. Photomicrographic studies of cross-sectioned coolant tubes seem to indicate that the effective thermal conductivity of the coatings is controlled by contact resistance between the particles, as a result of the spraying process, and not the thermal conductivity of the bulk materials.

Price, H. G., Jr.↗
NASA NTRS1992

Study of basic physical processes in liquid rocket engines

Inconsistencies between analytical results and measurements for liquid rocket thrust chamber performance, which escape suitable explanations, have motivated the examination of the basic phys ical modeling formulations as to their unlimited application. The publication of Prof. D. Straub's book, 'Thermofluid-dynamics of Optimized Rocket Propulsions,' further stimulated the interest of understanding the gas dynamic relationships in chemically reacting mixtures. A review of other concepts proposed by Falk-Ruppel (Gibbsian Thermodynamics), Straub (Alternative Theory, AT), Prigogine (Non-Equilibrium Thermodynamics), Boltzmann (Kinetic Theory), and Truesdell (Rational Mechanism) has been made to obtain a better understanding of the Navier-Stokes equation, which is now used extensively for chemically reacting flow treatment in combustion chambers. In addition to the study of the different concepts, two workshops were conducted to clarify some of the issues. The first workshop centered on Falk-Ruppel's new 'dynamics' concept, while the second one concentrated on Straub's AT. In this report brief summaries of the reviewed philosophies are presented and compared with the classical Navier-Stokes formulation in a tabular arrangement. Also the highlights of both workshops are addressed.

Wu, S. T.↗
NASA NTRS1959

Propellant Vaporization as a Criterion for Rocket-Engine Design; Experimental Performance, Vaporization and Heat-Transfer Rates with Various Propellant Combinations

Experimental combustion efficiencies of eleven propellant combinations were determined as a function of chamber length. Efficiencies were measured in terms of characteristic exhaust velocities at three chamber lengths and in terms of gas velocities. The data were obtained in a nominal 200-pound-thrust rocket engine. Injector and engine configurations were kept essentially the same to allow comparison of the performance. The data, except for those on hydrazine and ammonia-fluorine, agreed with predicted results based on the assumption that vaporization of the propellants determines the rate of combustion. Decomposition in the liquid phase may be.responsible for the anomalous behavior of hydrazine. Over-all heat-transfer rates were also measured for each combination. These rates were close to the values predicted by standard heat-transfer calculations except for the combinations using ammonia.

Clark, Bruce J.↗
NASA NTRS2002

A Model for Dynamic Simulation and Analysis of Tether Momentum Exchange

Momentum-exchange/electrodynamic reboost (MXER) tether systems may enable high-energy missions to the Moon, Mars, and beyond by serving as an 'upper stage in space'. Existing rockets that use an MXER tether station could double their capability to launch communications satellites and help improve US competitiveness. A MXER tether station would boost spacecraft from low Earth orbit to a high-energy orbit quickly, like a high-thrust rocket. Then, using the same principles that make an electric motor work, it would slowly rebuild its orbital momentum by pushing against the Earth's magnetic field-without using any propellant. One of the significant challenges in developing a momentum-exchange/electrodynamic reboost tether systems is in the analysis and design of the capture mechanism and its effects on the overall dynamics of the system. This paper will present a model for a momentum-exchange tether system that can simulate and evaluate the performance and requirements of such a system.

Canfield, Stephen↗
NASA NTRS2024

Residual Thrust in a Solid Rocket Motor: A Model Applied to the Mars Ascent Vehicle

A model for solid rocket motor residual thrust is built from the concept that hot nozzle phenolic or carbon materials provide a radiating source of heat flux to the insulation, leading to continued off-gassing by pyrolysis. This model is 1. Fitted to flight datasets and scaled to the Mars Ascent Vehicle second stage motor (MAV SRM2), and 2. Derived with MAV SRM2 design and material estimates. The two methods show a potential range of modeling uncertainty for use by mission analysts. Further loss methods that could reduce the maximum residual impulse are identified but left to further work to evaluate.

Solid rocket motors↗
NASA NTRSDate not supplied

Residual Thrust in a Solid Rocket Motor: A Model Applied to the Mars Ascent Vehicle

A model for solid rocket motor residual thrust is built from the concept that hot nozzle phenolic or carbon materials provide a radiating source of heat flux to the insulation, leading to continued off-gassing by pyrolysis. This model is 1. Fitted to flight datasets and scaled to the Mars Ascent Vehicle second stage motor (MAV SRM2), and 2. Derived with MAV SRM2 design and material estimates. The two methods show a potential range of modeling uncertainty for use by mission analysts. Further loss methods that could reduce the maximum residual impulse are identified but left to further work to evaluate.

Solid rocket motors↗
NASA NTRS1993

Effect of hydrogen exposure on a Cu-8 Cr-4 Nb alloy

The advanced regeneratively cooled rocket thrust chamber may require new materials to achieve long life and improved performance. Current materials such as NARloy-Z (Cu-3 wt. percent Ag-0.5 wt. percent Zr), while highly conductive, do not have sufficient high temperature strength and creep resistance to meet the projected needs of advanced rocket motors. A Cu-8 at. percent Cr-4 at. percent Nb (Cu-8 Cr-4 Nb) alloy has been identified as a promising material for this application. However, hydrogen embrittlement is a concern given the presence of high pressure, high temperature hydrogen in regeneratively cooled rocket motors. Thermodynamic analysis of the reaction between Cr-rich Cr2Nb and H2 showed that there is a possibility of reaction at temperatures up to 323 K in a 35 MPa H2 environment. Above 323 K the pressure necessary to achieve reaction rapidly increased beyond the range experienced in rocket motors. Tensile specimens exposed in 34.5 MPa H2 at room temperatures and during cycling to 705 C did not show any degradation of properties. No evidence of reaction was observed for Cr2Nb precipitate observed on the fracture surfaces. Based on these results the Cu-8 Cr-4 Nb alloy was judged to be sufficiently stable for use in rocket motors.

David L Ellis↗
NASA NTRS1994

Analyzing Thermal Conditions In Rocket Engines

Computer code, RTE, developed to perform three-dimensional thermal analyses of rocket thrust chambers. Calculates rate of heat transfer from combustion gases to coolant, coolant-temperature rise and pressure drop, and temperature profiles within cooling-jacket wall. Also calculates combustion-gas wall static pressure, temperature and enthalpy, as well as coolant pressure, temperature, and mach number for all stations. Program used for any propellant combination and most coolants commonly used in rockets. Code used for both regeneratively and radiatively cooled engines. However, in case of regeneratively cooled engines, applicability limited to engines featuring single-pass cooling and rectangular cooling channels.

Naraghi, M. H. N.↗
NASA NTRS1961

Pressure Loads Produced on a Flat-Plate Wing By Rocket Jets Exhausting in a Spanwise Direction Below the Wing and Perpendicular to a Free-Stream Flow of Mach Number 2.0

An investigation at a Reynolds number per foot of 14.4 x 10(exp 6) was made to determine the pressure loads produced on a flat-plate wing by rocket jets exhausting in a spanwise direction beneath the wing and perpendicular to a free-stream flow of Mach number 2.0. The ranges of the variables involved were (1) nozzle types - one sonic (jet Mach number of 1.00), two supersonic (jet Mach numbers of 1.74 and 3.04),. and one two-dimensional supersonic (jet Mach number of 1.71); (2) vertical nozzle positions beneath the wing of 4, 8 and 12 nozzle-throat diameters; and (3) ratios of rocket-chamber total pressure to free- stream static pressure from 0 to 130. The incremental normal force due to jet interference on the wing varied from one to two times the rocket thrust and generally decreased as the pressure ratio increased. The chordwise coordinate of the incremental-normal-force center of pressure remained upstream of the nozzle center line for the nozzle positions and pressure ratios of the investigation. The chordwise coordinate approached zero as the jet vertical distance beneath the wing increased. In the spanwise direction there was little change due to varying rocket-jet position and pressure ratio. Some boundary-layer flow separation on the wing was observed for the rocket jets close to the wing and at the higher pressure ratios. The magnitude of the chordwise and spanwise pressure distributions due to jet interference was greatest for rocket jets close to the wing and decreased as the jet was displaced farther from the wing. The design procedure for the rockets used is given in the appendix.

Falanga, Ralph A.↗
NASA NTRS1980

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

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

Boretz, J. E.↗
NASA NTRS1993

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

Niiya, Karen E.↗