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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 397 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Results of a rocket-Nimbus sounder comparison experiment

Stratospheric temperature data obtained from instrumentation on board Nimbus 3 and 4 are compared with conventional rocket soundings to determine the compatibility of measurements. The Arcasonde 1A and Datasonde were the primary rocketsondes used, but acoustic grenade and pitot probe soundings were also involved. Observed temperature profiles and computed radiances obtained from rocket soundings are compared with 15-microns data provided by the satellite infrared spectrometers, the infrared interferometer spectrometer, and the selective chopper radiometer onboard the satellites. A reasonable agreement between the rocket and satellite data is indicated, but variations are noted which are related to the sounding techniques.

Miller, A. J.↗
NASA NTRS1971

Nozzle and cavity wall cooling limitations on specific impulse of a gas-core nuclear rocket.

Experimental and theoretical study of the performance of a gas-core nuclear rocket, showing that in a uranium plasma nuclear rocket with an 8-ft cavity diameter the cavity wall can be cooled up to a power level of 7400 MW when the rocket operates at a pressure of 1000 atm and the propellant mass flow rate is 10 lbm/sec. A maximum cavity specific impulse of 5800 sec was obtained under such operating conditions. The fact that the wall heat flux was much lower for reactor powers below this level is linked to the presence of a relatively cool, opaque insulating layer of seeded propellant between the hot plasma and the solid wall. The additional coolant required for nozzle protection reduced the maximum cavity specific impulse to 5200 sec.

Kascak, A. F.↗
NASA NTRS1971

Technology for low cost solid rocket boosters.

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

Ciepluch, C.↗
NASA NTRS1972

Development of boron epoxy rocket motor chambers.

A 71 cm diameter 74 cm length boron/epoxy composite rocket motor chamber was designed based on the geometric configuration of the JPL Applications Technology Satellite titanium alloy apogee motor chamber. Because analyses showed large stress concentrations in the domes, the configuration was modified using the same basic constraints for openings and attachments. The rocket motor chamber was then fabricated by filament winding with boron/epoxy tape and hydrostatically tested to failure at 264 N/sq cm, 57.2 N/sq cm above the design value. Two more rocket motor chambers were fabricated with the same basic constraints, but shortened to 57.6 cm for a smaller propellant load. The first of these short chambers failed in proof because of filament winding fabrication difficulties. The second chamber was successfully fabricated and passed the hydrostatic proof test.

Jensen, W. M.↗
NASA NTRS1972

Pulse performance analysis for small hypergolic propellant rocket engines

Small rocket engine tests were conducted for the purpose of obtaining pulse performance data to aid in preliminary design and evaluation of attitude control systems. Both monopropellant and hypergolic bipropellant engines of thrust levels from 5 to 445 N (1 to 100 lb) were tested. The performance data for the hypergolic propellant rockets are compared with theoretical performance calculated from idealized chamber filling and evacuation characteristics. Electromechanical delays in valve response and heat transfer characteristics were found to cause substantial deviation between theoretical performance and test performance. The theoretical analysis is modified to obtain a semiempirical model for hypergolic propellant rockets.

Smith, G. W.↗
NASA NTRS1972

A program of high resolution X-ray astronomy using sounding rockets

Two Aerobee 170 sounding rocket payloads were flown at the White Sands Missile Range: (1) a focusing X-ray collector on 31 March 1972; and (2) a high resolution telescope on 4 August 1972. Data has been reduced from each of these flights. In the first flight both the rocket and the experiment instrumentation performed adequately, and it is clear that at least the minimum scientific objectives were attained. In the second flight the attitude control system failed to point the telescope at the target for a sufficient length of time. However examination of final preflight checkout data and some flight data indicate that the instrumentation for this rocket payload was functioning according to expectations.

Source record↗
NASA NTRS1972

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

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

Source record↗
NASA NTRS1972

Liquid propellant rockets.

A brief overview of the state of knowledge in liquid rocket technology is presented and examples are provided of instances where some fundamental principles of chemistry, fluid mechanics, and mathematics can be applied. A liquid propellant rocket classification is discussed together with rocket system performance, applications for liquid propellants, the effective exhaust velocity, aspects of simplified nozzle expansion, questions about theoretical propellant performance, the effect of chamber pressure on equilibrium performance, and the kinetic recombination in nozzles. Details of propellant combustion are examined, giving attention to propellant injection, evaporation-controlled combustion, combustion instability, and monopropellant decomposition.

Dipprey, D. F.↗
NASA NTRS1973

Comparison of particle results from a sounding rocket and a geosynchronous satellite

Auroral data from the ATS 5 geosynchronous satellite and an auroral rocket are compared with the purpose of illustrating some problems which will arise in comparison of data from the GEOS satellite and coordinated rocket experiments. It is recommended that all sky cameras be used in such studies and that since conjugacy between satellite and rocket cannot be guaranteed, the difference should be measured.

Johnstone, A. D.↗
NASA NTRS1973

The resolution of temporal and spatial variations in auroral particle measurements from sounding rockets

The problems involved in the resolution of the temporal/spatial ambiguity in sounding rocket observations are discussed with reference to an experiment involving a mother-daughter rocket combination, together with a flashing light beacon on the rocket and television cameras on the ground. The ambiguity was successfully removed, enabling a proper interpretation of the observations of precipitating low energy electrons to be made.

Johnstone, A. D.↗
NASA NTRS1974

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

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

Baran, W. J.↗
NASA NTRS1974

Liquid rocket combustion computer model with distributed energy release. DER computer program documentation and user's guide, volume 1

A computer program for analyzing rocket engine performance was developed. The program is concerned with the formation, distribution, flow, and combustion of liquid sprays and combustion product gases in conventional rocket combustion chambers. The capabilities of the program to determine the combustion characteristics of the rocket engine are described. Sample data code sheets show the correct sequence and formats for variable values and include notes concerning options to bypass the input of certain data. A seperate list defines the variables and indicates their required dimensions.

Combs, L. P.↗
NASA NTRS1973

Linear rocket engine design-fabrication-testing

A new rocket engine concept that is presently being evaluated offers significant performance, cost, and configurational advantages. The Linear Rocket Engine System, (LRES) combines the unique performance and operational advantages of the aerospike nozzle with the configuration versatility of low cost modular combustors, which results in a high-performance rocket engine package designed specifically for use in advanced configuration vehicles. This paper describes the design, fabrication, and testing of the LRES concept to evaluate its potential for vehicle integration. The program was conducted under the sponsorship of NASA's Marshall Space Flight Center.

Fuller, P. N.↗
NASA NTRS1974

Transient desorption of water vapor - A potential source of error in upper atmosphere rocket experiments

Results of measurements of the outgassing rates of samples of materials and surface finishes used on the outer skins of rocket-borne experiment packages in simulated rocket ascents. The results showed outgassing rates for anodized aluminum in the second minute of flight which are two to three orders of magnitude higher than those given in typical tables of outgassing rates. The measured rates for aluminum with chromate conversion surface coatings were also abnormally high. These abnormally high initial rates fell quickly after about five to ten minutes to values comparable with those in the published literature. It is concluded that anodized and chromate conversion coatings on the aluminum outer surfaces of a sounding rocket experiment package will cause gross distortion of the true water vapor environment.

Kendall, B. R. F.↗
NASA NTRS1974

A rocket spectroscopic payload in support of the Apollo Telescope Mount experiments

The scientific instrumentation and other payload systems of a solar rocket experiment are described in detail. The objectives of the rocket payload were: (1) to carry out high-spectral-resolution measurements of a coronal active region in the X-ray and extreme ultraviolet regions at the same time as high-spatial-resolution measurements were being made of the same active region by the Apollo Telescope Mount experiments flown on Skylab; and (2) to derive a physical model of the conditions in the coronal active regions, which dominate the X-ray spectrum of the nonflaring active sun, on the basis of data obtained from both the rocket instrumentation and several of the Apollo Telescope Mount experiments.

Rugge, H. R.↗
NASA NTRS1974

NASA plans for space-processing experiments on sounding rockets

Planning for sounding rocket space processing experiments in the second half of 1975 is reported. Currently operational sounding rocket systems are capable of carrying payloads containing more than 50 kg of experimental apparatus on trajectories including between 5 and 10 min of coasting flight above 100 km altitude. When equipped with stabilization systems to nullify angular velocities in their rotational degrees of freedom, these vehicles can provide an environment in which accelerations are less than 1 mm/sec per 2. The cost of flight operations to implement a rocket experiment program will be of the order of 2000 dollars per kg of discretionary payload, and within the next two years newly operational systems are expected to reduce the cost to about 700 dollars per kg.

Bredt, J. H.↗
NASA NTRS1975

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

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.

Stewart, R. B.↗
NASA NTRS1974

Near-simultaneous measurement of low-energy electrons by sounding rocket and satellite

An auroral sounding rocket was launched from Fort Churchill at 0453 UT on April 24, 1968. At the time of the rocket flight, the Vela 3B satellite was passing through the magnetotail plasma sheet at very nearly the same magnetic local time. Electron spectra obtained from experiments on these vehicles are well described by central Maxwell-Boltzmann populations, with indications of a high-energy tail; the satellite spectra also suggest the presence of a low-energy tail. The higher-temperature spectra measured on the rocket may be derived from the satellite spectra by adiabatic compression.

Rearwin, S.↗