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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
DOE OSTI2023

An Architectural Survey of the Area 25 Nuclear Rocket Development Station, Nevada National Security Site, Nye County, Nevada, Volume 3 Of 3

The purpose of the current architectural survey is to record the Nuclear Rocket Development Station (NRDS) district, which is recommended eligible for listing to the NRHP under all four criteria as the testing center for the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) program from the initial boundary survey in 1956 to the end of nuclear rocket development on January 5, 1973. Resources in the NRDS district were evaluated to determine if they are eligible for listing to the National Register as district contributors. The current survey identified the Reactor Control Point subdistrict to provide a temporal, spatial, and functional framework for organizing the resources involved in the NRDS operations. However, the subdistrict was not evaluated as a potential historic district against the Secretary of Interior’s Significance Criteria.

54 ENVIRONMENTAL SCIENCES↗
DOE OSTI2018

Nonadiabatic Flamelet Formulation for Predicting Wall Heat Transfer in Rocket Engines

A flamelet-based combustion model is introduced for the prediction of wall heat transfer in rocket engines and confined combustion systems. To account for convective heat loss due to the interaction of the flame with the wall, a permeable thermal boundary condition is introduced in the counterflow diffusion flame configuration. The solution of the resulting nonadiabatic flame structure forms a three-dimensional manifold, which is parameterized in terms of mixture fraction, progress variable, and temperature. The performance of the model is first assessed through a direct numerical simulation analysis of an H 2 /O 2 diffusion flame that is stabilized at an inert isothermal wall. The developed nonadiabatic flamelet model is shown to accurately predict the temperature, chemical composition, and wall heat transfer. Combined with a presumed probability density function closure, the model is then applied to large-eddy simulation of a single-injector rocket combustor to examine effects of heat transfer on the turbulent flame structure in rocket engines.

42 ENGINEERING↗
DOE OSTI2024

Satellite-Based Assessment of Rocket Launch and Coastal Change Impacts on Cape Canaveral Barrier Island, Florida, USA

The Cape Canaveral Barrier Island, home to the National Aeronautics and Space Administration (NASA)’s Kennedy Space Center and the United States (U.S.) Space Force’s Cape Canaveral Space Force Station, is situated in a unique ecological transition zone that supports diverse wildlife. This study evaluates the recent changes in vegetation cover (2016–2023) and dune elevation (2007–2017) within the Cape Canaveral Barrier Island using high-resolution optical satellite and light detection and ranging (LiDAR) data. The study period was chosen to depict the time period of a recent increase in rocket launches. The study objectives include assessing changes in vegetation communities, identifying detectable impacts of liquid propellant launches on nearby vegetation, and evaluating dune elevation and tide level shifts near launchpads. The results indicate vegetation cover changes, including mangrove expansion in wetland areas and the conversion of coastal strands to denser scrubs and hardwood forests, which were likely influenced by mild winters and fire management. While detectable impacts of rocket launches on nearby vegetation were observed, they were less severe than those caused by solid rocket motors. Compounding challenges, such as rising tide levels, beach erosion, and wetland loss, potentially threaten the resilience of launch operations and the surrounding habitats. The volume and scale of launches continue to increase, and a balance between space exploration and ecological conservation is required in this biodiverse region. This study focuses on the assessment of barrier islands’ shorelines.

54 ENVIRONMENTAL SCIENCES↗
DOE OSTI2024

Simulations of the collection of mesospheric dust particles with a rocket instrument

We investigate the collection of dust particles in the mesosphere with the MESS (MEteoric Smoke Sampler) instrument that is designed to fly on a sounding rocket. We assume that the ice particles that form in the polar mesosphere between 80 and 85 km altitude in summer contain meteoric smoke particles; and these should be collected with MESS. The instrument consists of a collection device with an opening and closure mechanism, as well as an attached conic funnel which increases the sampling area in comparison to the collection area. Dust particles are collected either directly after passing through the instrument or indirectly after colliding with and fragmenting on the funnel wall. We calculate the dust and fragment trajectories in the detector to determine the collection efficiency for different particle sizes, rocket velocities, and heights, and we find the final velocities and the temperatures of the particles. The considered design has a sampling area of 62.78 mm diameter and a collection area of 20 mm diameter. For the conditions at the rocket launch site in Andøya, Norway, we estimate the collection of meteoric smoke particles contained in the ice particles to be ~10 12 –10 14 amu mm –2 . The estimated temperatures suggest that the composition of these smoke particles is not affected by the collection. Our calculations also show that keeping the instrument open above 85 km altitude increases the amount of small smoke particles that are directly collected. The directly collected smoke particles are heated as they decelerate, which can affect their composition.

47 OTHER INSTRUMENTATION↗
NASA NTRS1949

Experimental Study of Gas-Flow Separation in Overexpanded Exhaust Nozzles for Rocket Motors

In a rocket motor having a de Laval nozzle with a fixed expansion ratio, there is only one pressure ratio (p(sub c)/p(sub o)) for which this expansion ratio is correct. At any other pressure ratio the gases will be either underexpanded or overexpanded, depending on whether the pressure ratio is higher or lower than the correct value. Rocket motor flame patterns illustrating the four possible regimes are shown in Figures 1 to 4, inclusive. In Figure 1, the pressure ratio is higher than the correct one for the fixed expansion ratio of the nozzle, resulting in underexpansion of the gases. in Figure 2, the pressure ratio is the correct one. In Figure 3, the pressure ratio is only slightly lower than the correct one, resulting in overexpansion without separation of the gases from the wall. In Figure 4, the pressure ratio is considerably below the correct one, resulting in overexpansion with separation of the gases from the nozzle wall. The present report is concerned primarily with this fourth type of flow, that is, overexpansion of the gases with separation of the flow from the wall. Such a situation exists in a rocket motor operating at a considerably lower altitude than that for which its exhaust nozzle was designed.

Cowles, Frederick B.↗
NASA NTRS1968

Description and Flight Performance Results of the WASP Sounding Rocket

A general description of the design and construction of the WASP sounding rocket and of the performance of its first flight are presented. The purpose of the flight test was to place the 862-pound (391-kg) spacecraft above 250 000 feet (76.25 km) on free-fall trajectory for at least 6 minutes in order to study the effect of "weightlessness" on a slosh dynamics experiment. The WASP sounding rocket fulfilled its intended mission requirements. The sounding rocket approximately followed a nominal trajectory. The payload was in free fall above 250 000 feet (76.25 km) for 6.5 minutes and reached an apogee altitude of 134 nautical miles (248 km). Flight data including velocity, altitude, acceleration, roll rate, and angle of attack are discussed and compared to nominal performance calculations. The effect of residual burning of the second stage motor is analyzed. The flight vibration environment is presented and analyzed, including root mean square (RMS) and power spectral density analysis.

De Pauw, J. F.↗
NASA NTRS1971

The United States sounding rocket program

The United States sounding rocket program is discussed. The program is concerned with the fields of solar physics, galactic astronomy, fields and particles, ionospheric physics, aeronomy, and meteorology. Sounding rockets are described with respect to propulsion systems, gross weight, and capabilities. Instruments used to conduct ionospheric probing missions are examined. Results of previously conducted sounding rocket missions are included.

Source record↗
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 1 to 100 lbs 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 and test performance. The theoretical analysis is modified to obtain a semi-empirical model for hypergolic propellant rockets which is demonstrated to be reasonably accurate for two different engine configurations over a considerable range of duty cycles.

Smith, G. W.↗
NASA NTRS1972

The geometry and physical properties of exhaust clouds generated during the static firing of S-1C and S-2 rocket engines

An experimental program was conducted during the static firing of the S-1C stage 13, 14, and 15 rocket engines and the S-2 stage 13, 14, and 15 rocket engines. The data compiled during the experimental program consisted of photographic recordings of the time-dependent growth and diffusion of the exhaust clouds, the collection of meteorological data in the ambient atmosphere, and the acquisition of data on the physical structure of the exhaust clouds which were obtained by flying instrumented aircraft through the clouds. A new technique was developed to verify the previous measurements of evaporation and entrainment of blast deflector cooling water into the cloud. The results of the experimental program indicate that at the lower altitudes the rocket exhaust cloud or plume closely resembles a free-jet type of flow. At the upper altitudes, where the cloud is approaching an equilibrium condition, structure is very similar to a natural cumulus cloud.

Forbes, R. E.↗
NASA NTRS1972

Liquid rocket performance computer model with distributed energy release

Development of a computer program for analyzing the effects of bipropellant spray combustion processes on liquid rocket performance is described and discussed. The distributed energy release (DER) computer program was designed to become part of the JANNAF liquid rocket performance evaluation methodology and to account for performance losses associated with the propellant combustion processes, e.g., incomplete spray gasification, imperfect mixing between sprays and their reacting vapors, residual mixture ratio striations in the flow, and two-phase flow effects. The DER computer program begins by initializing the combustion field at the injection end of a conventional liquid rocket engine, based on injector and chamber design detail, and on propellant and combustion gas properties. It analyzes bipropellant combustion, proceeding stepwise down the chamber from those initial conditions through the nozzle throat.

Combs, L. P.↗
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.↗