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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Space Shuttle reaction control subsystem propellant tank masked screen test

During final development testing of the Space Shuttle Reaction Control Subsystem propellant tanks, a problem with pressure transients in the system was uncovered. Due to the nature of the tanks, performance tests to determine the impact of the transients on the expulsion efficiency of the tanks could not directly simulate the actual conditions which would be present in a low-gravity environment. However, by masking or covering various segments of the propellant acquisition device, a good simulation of a low-gravity environment was achieved in ground testing.

Henderson, J. B.↗
NASA NTRS1972

Evacuation areas for transportation accidents involving propellant tank pressure bursts

Evacuation areas are defined for those transportation accidents where volatile chemical propellant tanks are exposed to fire in the wreckage and eventually explode with consequent risks from fragments in surrounding populated areas. An evacuation area with a minimum radius of 600 m (2000 ft) is recommended to limit the statistical probability of fatality to one in 100 such accidents. The result was made possible by the derivation of a distribution function of distances reached by fragments from bursting chemical car tanks. Data concerning fragments was obtained from reports or tank car pressure bursts between 1958 and 1971.

Siewert, R. D.↗
NASA NTRS1972

Evacuation areas for transportation accidents involving propellant tank pressure bursts

Evacuation areas are defined for those transportation accidents where volatile chemical propellant tanks are exposed to fire in the wreckage and eventually explode with consequent risks from fragments in surrounding populated areas. An evacuation area with a minimum radius of 600 m is recommended to limit the statistical probability of fatality to one in 100 such accidents. The result of this study was made possible by the derivation of a distribution function of distances reached by fragments from bursting chemical car tanks. Data concerning fragments were obtained from reports of tank car pressure bursts between 1958 and 1971.

Siewert, R. D.↗
NASA NTRS2011

Sloshing in the Liquid Hydrogen and Liquid Oxygen Propellant Tanks After Main Engine Cut Off

NASA Marshall Space Flight Center is designing and developing the Main Propulsion System (MPS) for Ares launch vehicles. Propellant sloshing in the liquid hydrogen (LH2) and liquid oxygen (LO2) propellant tanks after Main Engine Cut Off (MECO) was modeled using the Volume of Fluid (VOF) module of the computational fluid dynamics code, CFD-ACE+. The present simulation shows that there is substantial sloshing side forces acting on the LH2 tank during the deceleration of the vehicle after MECO. The LH2 tank features a side wall drain pipe. The side loads result from the residual propellant mass motion in the LH2 tank which is initiated by the stop of flow into the drain pipe at MECO. The simulations show that radial force on the LH2 tank wall is less than 50 lbf and the radial moment calculated based up through the center of gravity of the vehicle is predicted to be as high as 300 lbf-ft. The LO2 tank features a bottom dome drain system and is equipped with sloshing baffles. The remaining LO2 in the tank slowly forms a liquid column along the centerline of tank under the zero gravity environments. The radial force on the LO2 tank wall is predicted to be less than 100 lbf. The radial moment calculated based on the center of gravity of the vehicle is predicted as high as 4500 lbf-ft just before MECO and dropped down to near zero after propellant draining stopped completely.

Kim, Sura↗
NASA NTRS1984

Evaluation of propellant tank insulation concepts for low-thrust chemical propulsion systems

An analytical evaluation of cryogenic propellant tank insulations for liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) was conducted. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three Shuttle-launched LTPS designed for Shuttle bay packaged payload densities of 56 kg/cu m, 40 kg/cu m and 24 kg/cu m. Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because Orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. It was noted this cost could be reduced significantly without increasing program risk.

Kramer, T.↗
NASA NTRSDate not supplied

Loci/STREAM Sharp Interface Model Validation of Cryogenic Propellant Tank Self-Pressurization

This paper documents a model validation effort for the prediction of cryogenic propellant tank selfpressurization physics using the Loci/STREAM sharp interface CFD model developed at NASA MSFC. Cryogenic fluid management (CFM) applications for long duration spaceflight missions require accurate modeling of heat transfer and thermodynamics. Therefore, detailed validation of those physics is important to ensure simulation accuracy. In the past, the MSFC Propulsion Fluid Dynamics Branch has validated and used the Loci/STREAM Volume of Fluid (VOF) module extensively for propellant slosh applications, which are dominated by fluid motion. For cases where the interface gas/liquid interface is essentially static, the Loci/STREAM sharp interface was developed for more rapid and efficient CFM analysis. In the present study, Loci/STREAM sharp interface simulations were compared with the NASA K-site self-pressurization experiment of a flight weight, partially full, LH2 tank in normal gravity. This study investigated several parametric effects and made a number of important observations on the performance of the Loci/STREAM sharp interface method: - First, a homogeneous model was derived to verify the CFD model. The homogeneous model assumes both gas and liquid phases are uniform and are at saturation temperature. The predicted pressure rise rates at 3.5 W/m2 and 2.0 W/m2 with a 49% fill level are all lower than the experimental measurement. This is expected due to the observed temperature stratification near the gas-liquid interface during testing. - The k-w and k-e turbulence models predict comparable pressurization rates, which are much lower than the experiment measurement. On the other hand, the pressure values and pressurization rates from both models are almost the same as those of the homogeneous model. The current turbulence model is too diffusive and is not capable of capturing the thermal stratification in the liquid and ullage. - When using the laminar flow model, the pressurization rate is found to be very close to the experimental value when a quasi-steady flow field is allowed to develop by continuously venting the tank prior to pressurization as was done during testing. On the other hand, without initial venting, the rate is not constant. Additionally, the development time to reach a quasi-steady flow field differed from the experiment due to the lack of turbulent mixing away from the gas-liquid interface that is expected in reality. - Using the laminar flow model and appropriate initial conditioning, the pressurization rate is within 15% of the experimental value for 49% fill level at 3.5 W/m 2 and is within 18% for 2.0 W/m 2 heat loading. - Using the laminar flow model and appropriate initial conditioning, the pressurization rate is within 18% of the experimental value for a 29% fill level and within 15% for an 83% fill level.

H. Q. Yang↗
NASA NTRS1962

Analysis of Thermal-Protection Systems for Space-Vehicle Cryogenic-Propellant Tanks

Analytical techniques are presented that permit the calculation of heat-transfer rates with various thermal-protection systems for liquid-cryogenic-propellant tanks subjected to on-board, solar, and planetary heat fluxes. The thermal-protection systems considered include using closely spaced reflective surfaces (foils) and widely spaced reflective surfaces (shadow shields), insulation, arrangement of vehicle components, orientation with respect to radiant heating sources, and coatings for the control of solar absorptivity. The effectiveness of these thermal-protection systems in reducing propellant heating is shown both for ideal heat-transfer models and for a simplified hydrogen-oxygen terminal stage on a Mars mission. The proper orientation of a space-vehicle cryogenic tank with respect to the Sun is one of the more beneficial methods of reducing the heating effect of solar flux. Shadow shields can be extremely effective in reducing the propellant heating due to both solar and on-board fluxes. However, low-altitude planet orbits can result in high propellant heating rates due to planetary radiation reflected from the shields. For low-altitude orbits of more than a few days, foils appear to be desirable for all cryogenic-tank surfaces. Foils are also effective in reducing the on-board heating. A choice of shadow shields or foils cannot be made until a particular vehicle and a particular mission are chosen. The thermal conductivity of insulation materials would have to be lower by about two orders of magnitude with no increase in density before insulation could compete with reflective surfaces for use in long-duration thermal protection of cryogenic tanks in space. To demonstrate the application of the methods devised, thermal-protection systems are developed for a hydrogen-oxygen terminal stage for typical Mars missions.

SPACE VEHICLE↗
NASA NTRSDate not supplied

Validation of Cryogenic Propellant Tank Self-Pressurization by Leveraging Reduced Order Modeling within Computational Fluid Dynamics Simulation

Validation of cryogenic propellant tank self-pressurization was performed using a hybrid Computational Fluid Dynamics (CFD) and reduced order modeling methodology. Data from a liquid hydrogen ground test conducted at the K-site facility at the National Aeronautics and Space Administration (NASA) Glenn Research Center was used for the validation effort. Liquid phase dynamics were explicitly resolved with a CFD tool. Vapor phase dynamics were modeled as a point mass that communicated heat from the tank wall to the liquid phase via a boundary condition used at the gas-liquid interface. The method proved to be more accurate, robust, and efficient than explicit resolution of the dynamics using a standard Volume of Fluid (VOF) methodology. The subject pressurization process was found to be heavily dependent upon both the relatively high liquid temperature gradient near the gas-liquid interface and the natural convection flow path. Modeling the gas-liquid interface as an immovable surface eliminated temperature gradient destroying gas-liquid interface velocities observed in VOF simulations, and correspondingly enabled more rapid simulation since interface advection was not allowed. The single phase computational domain also facilitated the ability to demonstrate spatial resolution convergence of natural convection cells within the liquid which significantly impacted the tank pressurization rate. This work was used to demonstrate the critical physics for tank self-pressurization and numerical methodologies that may be used to best resolve those physics. The findings informed development and operation of production level CFD tools used in the Fluid Dynamics Branch at NASA Marshall Space Flight Center.

J. M. Brodnick↗
NASA NTRS1962

Analysis of Thermal-Protection Systems for Space-Vehicle Cryogenic-Propellant Tanks

Analytical techniques are presented that permit the calculation of heat-transfer rates with various thermal-protection systems for liquid-cryogenic-propellant tanks subjected to on-board, solar, and planetary heat fluxes . The effectiveness of these protection systems in reducing propellant heating is shown both for ideal heat-transfer models and for a simplified hydrogen-oxygen terminal stage used for typical Mars missions.

Smolak, George R.↗
NASA NTRS1971

An experimental study of axisymmetric modes in various propellant tanks containing liquid

An experimental study of several models used to analyze the longitudinal dynamics of oscillations in launch vehicle propellant tanks, is presented. Several vehicle structural configurations are considered: hemispherical and spherical, cylindrical and spherical, and semiellipsoidal and ellipsoidal. Various bulkheads and flanges are used with each configuration. The results are given in graphs.

Kana, D. D.↗
NASA NTRS1965

Cryogenic Storage of Helium for Propellant Tank Pressurization

This paper has been prepared in response to many inquiries regarding the application of cryogenic helium storage to propellant tank pressurization systems. The high weight penalties associated with conventional helium storage systems have prompted the development of a Supercritical Helium Storage and Supply System for this application.

Tyler, J. S.↗