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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.

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At least 901 records · Page 50
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Miniature Rocket Motor for Aircraft Stall/Spin Recovery

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

Lucy, M. H.↗
NASA NTRS1985

Structural and mechanical design challenges of space shuttle solid rocket boosters separation and recovery subsystems

The design of the space shuttle solid rocket booster (SRB) subsystems for reuse posed some unique and challenging design considerations. The separation of the SRBs from the cluster (orbiter and external tank) at 150,000 ft when the orbiter engines are running at full thrust meant the two SRBs had to have positive separation forces pushing them away. At the same instant, the large attachments that had reacted launch loads of 7.5 million pounds thrust had to be servered. These design considerations dictated the design requirements for the pyrotechnics and separation rocket motors. The recovery and reuse of the two SRBs meant they had to be safely lowered to the ocean, remain afloat, and be owed back to shore. In general, both the pyrotechnic and recovery subsystems have met or exceeded design requirements. In twelve vehicles, there has only been one instance where the pyrotechnic system has failed to function properly.

Woodis, W. R.↗
NASA NTRS1984

Atomization data requirements for rocket combustor modeling

The complex computer codes, which model liquid rocket combustors, require information about the distribution and atomization of these liquid reactants. The available information is, in general, of questionable validity and applicability. Authors and users of combustion codes are often unaware of, or underestimate the importance of, these deficiencies in atomization data. These deficiencies and their importance are examined. Results of analyses performed with a state-of-the-art rocket combustion code are presented which demonstrate the important effects of such atomization information as initial droplet sizes and size distribution on vaporization rate and losses. Also, the questionable aspects and inapplicability of the available atomization data are discussed. One important and often neglected or misunderstood aspect of atomization data is the differences between spatial (concentration) and flux (often called temporal) droplet size distributions. These are described, and a computer model constructed to assess the difference between concentration and flux droplet size distributions is described and results presented. Experimental data are also given to demonstrate this difference. Finally, experimental results are presented that demonstrate the very great, and often neglected effect, of the local gas velocity field on atomization.

Ferrenberg, A. J.↗
NASA NTRS1985

Rocket thrust chamber thermal barrier coatings

Subscale rocket thrust chamber tests were conducted to evaluate the effectiveness and durability of thin yttria stabilized zirconium oxide coatings applied to the thrust chamber hot-gas side wall. The fabrication consisted of arc plasma spraying the ceramic coating and bond coat onto a mandrell and then electrodepositing the copper thrust chamber wall around the coating. Chambers were fabricated with coatings .008, and .005 and .003 inches thick. The chambers were thermally cycled at a chamber pressure of 600 psia using oxygen-hydrogen as propellants and liquid hydrogen as the coolant. The thicker coatings tended to delaminate, early in the cyclic testing, down to a uniform sublayer which remained well adhered during the remaining cycles. Two chambers with .003 inch coatings were subjected to 1500 thermal cycles with no coating loss in the throat region, which represents a tenfold increase in life over identical chambers having no coatings. An analysis is presented which shows that the heat lost to the coolant due to the coating, in a rocket thrust chamber design having a coating only in the throat region, can be recovered by adding only one inch to the combustion chamber length.

Quentmeyer, R. J.↗
NASA NTRS1985

Liquid rocket combustor computer code development

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

Liang, P. Y.↗
NASA NTRS1985

Artificial plasma density structures produced by energetic electron beams from rockets and spacecraft

Recent rocket and Space Shuttle experiments have demonstrated the capability to launch electron beams of moderate power (100 W to 10 kW) into the earth's ionosphere and magnetosphere. This letter describes how such beams, when fired from rockets or satellites, can create significant ionization in the E- and F-regions of the ionosphere. Through proper selection of beam-related parameters, an interesting variety of plasma density structures, including plasma sheets and plasma filaments, can be created and studied over periods of 30 minutes to 1 hour, depending on the rate of plasma recombination and the density of the ambient plasma. Observations of these structures can give new information relating to the physics of plasma density structures in the ionosphere, and the effects these features have upon the scattering of radio waves. It is also possible that observations of the density structures will provide a new means for studying neutral winds and electrodynamic phenomena in the ionosphere.

Banks, P. M.↗
NASA NTRS1986

Solar Thermal Rocket Propulsion

Paper analyzes potential of solar thermal rockets as means of propulsion for planetary spacecraft. Solar thermal rocket uses concentrated Sunlight to heat working fluid expelled through nozzle to produce thrust.

Sercel, J. C.↗
NASA NTRS1985

Analytical investigation of solid rocket nozzle failure

On April 5, 1983, an Inertial Upper Stage (IUS) spacecraft experienced loss of control during the burn of the second of two solid rocket motors. The anomaly investigation showed the cause to be a malfunction of the solid rocket motor. This paper presents a description of the IUS system, a failure analysis summary, an account of the thermal testing and computer modeling done at Marshall Space Flight Center, a comparison of analysis results with thermal data obtained from motor static tests, and describes some of the design enhancement incorporated to prevent recurrence of the anomaly.

Mccoy, K. E.↗
NASA NTRS1985

Rocket thrust chamber thermal barrier coatings

A research program was conducted to generate data and develop analytical techniques to predict the performance and reliability of ceramic thermal barrier coatings in high heat flux environments. A finite element model was used to analyze the thermomechanical behavior of coating systems in rocket thrust chambers. Candidate coating systems (using a copper substrate, NiCrAlY bond coat and ZrO2.8Y2O3 ceramic overcoat) were selected for detailed study based on photomicrographic evaluations of experimental test specimens. The effects of plasma spray application parameters on the material properties of these coatings were measured and the effects on coating performance evaluated using the finite element model. Coating design curves which define acceptable operating envelopes for seleted coating systems were constructed based on temperature and strain limitations. Spray gun power levels was found to have the most significant effect on coating structure. Three coating systems were selected for study using different power levels. Thermal conductivity, strain tolerance, density, and residual stress were measured for these coatings. Analyses indicated that extremely thin coatings ( 0.02 mm) are required to accommodate the high heat flux of a rocket thrust chamber and ensure structural integrity.

Batakis, A. P.↗
NASA NTRS1986

Rocket and laboratory studies in aeronomy and astrophysics

Progress from March 1, 1986 to August 31, 1986 is covered and includes the work performed in response to a proposal entitled A Spartan Payload for Spatially Resolved Spectroscopy of Extended Faint Sources in the Extreme Ultraviolet (EUV). During this period, one rocket was launched, the reflight of the payload to observe Halley's comet on March 13, 1986. Most of the effort during this period was concentrated on detailed mechanical and electronic design of a Spartan payload and on the reduction and analysis of the data from the two Halley rocket flights and from the UVX experiment which flew on STS-61C in January 1986.

Feldman, P. D.↗
NASA NTRS1986

Bipropellant rocket exhaust plume analysis on the Galileo spacecraft

This paper describes efforts to quantify the contaminant flow field produced by 10 N thrust bipropellant rocket engines used on the Galileo spacecraft. The prediction of the composition of the rocket exhaust by conventional techniques is found to be inadequate to explain experimental observations of contaminant deposition on moderately cold (200 K) surfaces. It is hypothesized that low volatility contaminants are formed by chemical reactions which occur on the surfaces. The flow field calculations performed using the direct simulation Monte Carlo method give the expected result that the use of line-of-sight plume shields may have very little effect on the flux of vapor phase contaminant species to a surface, especially if the plume shields are located so close to the engine that the interaction of the plume with the shield is in the transition flow regime. It is shown that significant variations in the exhaust plume composition caused by nonequilibrium effects in the flow field lead to very low concentrations of species which have high molecular weights in the more rarefied regions of the flow field. Recommendations for the design of spacecraft plume shields and further work are made.

Guernsey, C. S.↗
NASA NTRS1987

Coal-Fired Rocket Engine

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

Anderson, Floyd A.↗
NASA NTRS1987

Ignition overpressure study from solid rocket motor firings

The objective was to investigate, through experimental means, the basic mechanisms influencing ignition overpressure and to determine ways to suppress ignition overpressure. Ignition overpressure was studied using solid rocket motors with geometry scaled at 1 percent of the Shuttle's Solid Rocket Boosters. Both water injection and aerosol foam were examined as a mean of reducing ignition overpressure. The results of the water injection tests indicate that a relatively small amount of water is sufficient to provide significant suppression. Of the flow rates tested, the lower water injection flow rates provided the best reduction of the ignition overpressure wave. Also, the test results show there is an optimum water flow rate range that provides the best suppression, and as this range is exceeded the effectiveness of water to reduce ignition overpressure is decreased. Aerosol foam provided very little reduction of ignition overpressure, but only small volumes of foam were used and further testing is necessary to determine its total effectiveness as a means of suppression.

Counter, Douglas D.↗
NASA NTRS1986

Application of boost guidance to NASA sounding rocket launch operations at the White Sands Missile Range

This paper addresses the unique problems associated with launching the Black Brant V, VIII, and IX sounding rocket vehicles at White Sands Missile Range (WSMR) and the significance of the introduction of the S19 to the NASA Goddard Space Flight Center Wallops Flight Facility sounding rocket program in terms of launch flexibility, improved impact dispersion, higher flight reliability, and reduced program costs. This paper also discusses salient flight results from NASA 36.011UL (the first S19 guided Black Brant launched at WSMR) and the NASA Comet Halley missions (36.010DL and 36.017DL).

Montag, W. H.↗
NASA NTRS1986

Mutual downwash interference effects on the normal force and roll driving coefficients of multi-stage rockets

For several years, the TAD2 computer program has been used to calculate the aerodynamic properties of sounding rockets. This program was limited, however, in that a maximum of two fin sets could be accommodated analytically. In addition, the roll moment coefficients did not consider downwash effects. Techniques are developed and equations derived for the expansion of capability to include three fin sets. The effects of downwash on the roll moment coefficients are also treated analytically for two and three stage rockets.

Ward, P. R.↗
NASA NTRS1986

Development of a reclosable clamshell nose fairing for sounding rocket NASA 21.003UI

A reclosable clamshell noise fairing system was developed and successfully flown on a Black Brant VC sounding rocket. The clamshell system involved the design of unique structural and mechanical subsystems to meet the experimenters requirements, withstand environmental conditions of sounding rocket flight and recovery, and be relatively easy to maintain with a clean interface between the experiment detectors and instrumentation section of the payload. The development of the equations of motion of the mechanism are illustrated and results of the test of the Engineering Test Model (ETM) are shown.

Ferragut, N. J.↗
NASA NTRS1986

Flip-Flop Recovery System for sounding rocket payloads

The design, development, and testing of the Flip-Flop Recovery System, which protects sensitive forward-mounted instruments from ground impact during sounding rocket payload recovery operations, are discussed. The system was originally developed to reduce the impact damage to the expensive gold-plated forward-mounted spectrometers in two existing Taurus-Orion rocket payloads. The concept of the recovery system is simple: the payload is flipped over end-for-end at a predetermined time just after parachute deployment, thus minimizing the risk of damage to the sensitive forward portion of the payload from ground impact.

Flores, A., Jr.↗