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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Coolant-side heat-transfer rates for a hydrogen-oxygen rocket and a new technique for data correlation

An experimental investigation was conducted to determine the coolant-side, heat transfer coefficients for a liquid cooled, hydrogen-oxygen rocket thrust chamber. Heat transfer rates were determined from measurements of local hot gas wall temperature, local coolant temperature, and local coolant pressure. A correlation incorporating an integration technique for the transport properties needed near the pseudocritical temperature of liquid hydrogen gives a satisfactory prediction of hot gas wall temperatures.

Schacht, R. L.↗
NASA NTRS1991

An investigation of voids formation mechanisms and their effects on freeze and thaw processes of lithium and lithium fluoride

The mechanisms of void formation during the cooldown and freezing of lithium coolant within the primary loop of SP-100 type systems are investigated. These mechanisms are: (1) homogeneous nucleation; (2) heterogeneous nucleation; (3) normal segregation of helium gas dissolved in liquid lithium; and (4) shrinkage of lithium during freezing. To evaluate the void formation potential due to segregation, a numerical scheme that couples the freezing and mass diffusion processes in both the solid and liquid regions is developed. The results indicated that the formation of He bubbles is unlikely by either homogeneous or heterogeneous nucleation during the cooldown process. However, homogeneous nucleation of He bubbles following the segregation of dissolved He in liquid lithium ahead of the solid-liquid interface is likely to occur. Results also show that total volume of He void is insignificant when compared to that of shrinkage voids. In viewing this, the subsequent research focuses on the effects of shrinkage void forming during freezing of lithium on subsequent thaw processes are investigated using a numerical scheme that is based on a single (solid/liquid) cell approach. The cases of lithium-fluoride are also investigated to show the effect of larger volume shrinkage upon freezing on the freeze and thaw processes. Results show that a void forming at the wall appreciably reduces the solid-liquid interface velocity, during both freeze and thaw, and causes a substantial rise in the wall temperature during thaw. However, in the case of Li, the maximum wall temperature was much lower than the melting temperature of PWC-11, which is used as the structure material in the SP-100 system. Hence, it is included that a formation of hot spots is unlikely during the startup or restart of the SP-100 system.

El-Genk, Mohamed S.↗
NASA NTRS1991

Shrinkage void formation and its effect on freeze and thaw processes of lithium and lithium-fluoride for space applications

The effects of shrinkage void forming during freezing of lithium and lithium fluoride on subsequent thaw processes are investigated using a numerical scheme that is based on a single (solid/liquid) cell approach. Results show that a void forming at the wall appreciably reduces the solid-liquid interface velocity, during both freeze and thaw, and causes a substantial rise in the wall temperature during thaw. However, in the case of Li, the maximum wall temperature was much lower than the melting temperature of PWC-11, which is used as the structure material in the SP-100 system. Hence, it is concluded that a formation of hot spots is unlikely during the startup or restart of the SP-100 system.

Yang, Jae Y.↗
NASA NTRS1959

Details of Exact Low Prandtl Number Boundary-Layer Solutions for Forced and For Free Convection

A detailed report is given of exact (numerical) solutions of the laminar-boundary-layer equations for the Prandtl number range appropriate to liquid metals (0.003 to 0.03). Consideration is given to the following situations: (1) forced convection over a flat plate for the conditions of uniform wall temperature and uniform wall heat flux, and (2) free convection over an isothermal vertical plate. Tabulations of the new solutions are given in detail. Results are presented for the heat-transfer and shear-stress characteristics; temperature and velocity distributions are also shown. The heat-transfer results are correlated in terms of dimensionless parameters that vary only slightly over the entire liquid-metal range. Previous analytical and experimental work on low Prandtl number boundary layers is surveyed and compared with the new exact solutions.

Sparrow, E. M.↗
NASA NTRSDate not supplied

Validation of Cryogenic Propellant Tank Filling using Computational Fluid Dynamics Simulation

Validation of cryogenic propellant tank filling was performed using the Computational Fluid Dynamics (CFD) solver Loci/STREAM-VoF. The validation effort helped identify modeling methodologies that enable NASA to best support its partners in both launch pad and on-orbit filling operations. Data from liquid hydrogen ground tests filled via jet injection were used for the validation effort which include a sensitivity to initial tank wall temperature. Initially hot walls are expected to yield rapid evaporation and possibly boiling. A Volume of Fluid (VoF) methodology was used to capture the gas-liquid interface. Rapid breakup of the liquid jet was observed in simulation results as liquid evaporated and expanded. Inflowing liquid transitioned to a contiguous jet as tank temperatures decreased, gas pressure increased, and saturation conditions at the incoming liquid temperature were approached. The final phase of filling was distinguished by rapid gas pressure rise due to a higher rate of gas volume compression than condensation at the liquid surface. Key physics of propellant tank filling were captured in the computational predictions, and opportunities for added simulation robustness and efficiency in future modeling efforts were identified.

CFM↗
NASA NTRS1991

Thermal calibration and analysis of a Bridgman flight furnace

A combination of computer modeling and experimental measurements is proposed which permits the calibration of the temperature profile for a spacecraft flight furnace. The Fluent code based on the finite difference method uses the furnace set-point temperatures as boundary conditions to model a Bridgman directional solidification furnace. Experimental measurements are employed as a comparison for the numerical analysis, and the radial gradients and wall temperatures are selected when the calculated values are sufficiently close to those from measurements. The calculations are found to agree with experimental data, and some results of the modeling are given which relate to specific furnace parameters. The CFD approach facilitates the transposition of furnace calibration to microgravitational operation in addition to providing precise initial radial temperature gradients and furnace-wall temperatures.

Knuteson, D. J.↗
NASA NTRS1977

Liquid rocket engine self-cooled combustion chambers

Self-cooled combustion chambers are chambers in which the chamber wall temperature is controlled by methods other than fluid flow within the chamber wall supplied from an external source. In such chambers, adiabatic wall temperature may be controlled by use of upstream fluid components such as the injector or a film-coolant ring, or by internal flow of self-contained materials; e.g. pyrolysis gas flow in charring ablators, and the flow of infiltrated liquid metals in porous matrices. Five types of self-cooled chambers are considered in this monograph. The name identifying the chamber is indicative of the method (mechanism) by which the chamber is cooled, as follows: ablative; radiation cooled; internally regenerative (Interegen); heat sink; adiabatic wall. Except for the Interegen and heat sink concepts, each chamber type is discussed separately. A separate and final section of the monograph deals with heat transfer to the chamber wall and treats Stanton number evaluation, film cooling, and film-coolant injection techniques, since these subjects are common to all chamber types. Techniques for analysis of gas film cooling and liquid film cooling are presented.

Source record↗
NASA NTRS1993

Analysis of film cooling in rocket nozzles

This report summarizes the findings on the NASA contract NAG8-212, Task No. 3. The overall project consists of three tasks, all of which have been successfully completed. In addition, some supporting supplemental work, not required by the contract, has been performed and is documented herein. Task 1 involved the modification of the wall functions in the code FDNS (Finite Difference Navier-Stokes) to use a Reynolds Analogy-based method. This task was completed in August, 1992. Task 2 involved the verification of the code against experimentally available data. The data chosen for comparison was from an experiment involving the injection of helium from a wall jet. Results obtained in completing this task also show the sensitivity of the FDNS code to unknown conditions at the injection slot. This task was completed in September, 1992. Task 3 required the computation of the flow of hot exhaust gases through the P&W 40K subscale nozzle. Computations were performed both with and without film coolant injection. This task was completed in July, 1993. The FDNS program tends to overpredict heat fluxes, but, with suitable modeling of backside cooling, may give reasonable wall temperature predictions. For film cooling in the P&W 40K calorimeter subscale nozzle, the average wall temperature is reduced from 1750R to about 1050R by the film cooling. The average wall heat flux is reduced by a factor of 3.

Woodbury, Keith A.↗
NASA NTRS1972

Heat-transfer testing procedures in phase B shuttle studies with emphasis on phase change data improvement

The procedures used in the application of the phase change technique to the phase B shuttle configuration are discussed along with factors which may affect data accuracy. These factors include variation of thermal properties of phase change model material, sensitivity of measured heat transfer coefficients to the assumed value of the adiabatic to total temperature ratio, and wall temperature effects. These sensitivities are illustrated in sample calculations for a shuttle geometry. Factors which may affect the visual clarity and interpretation of phase change data are discussed, and a method of improving photographic data quality through the use of polarized light is presented.

Throckmorton, D. A.↗
NASA NTRS1974

Comments on Crocco's solution and the independence principle for compressible turbulent boundary layers

The purpose of the present note is to show that on a flat plate where both the wall temperature and mean wall pressure are constant, neither of the limitations of parallel flow or of unity for the turbulent Prandtl number are required in order for the Crocco solution to apply to the turbulent boundary-layer flow. It is shown herein that this result is subject to restrictions on the magnitude of pressure fluctuations. The same analysis is generalized to show that the compressible turbulent boundary layer on an isothermal swept flat plate is independent of the spanwise flow if the molecular Prandtl number is unity.

Beckwith, I. E.↗
NASA NTRS1976

Laminar free convection from a vertical cone with uniform surface heat flux

The laminar free convection flow from a right circular cone with prescribed uniform wall flux condition is studied. The governing boundary-layer equations are analyzed by the technique of similarity transformation. Numerical solutions to the transformed equations are given for Prandtl numbers 0.72, 1, 2, 4, 6, 8, 10, and 100. Expressions for both wall temperature distribution and wall skin friction distribution at Prandtl number tending to infinity are also presented.

Lin, F. N.↗
NASA NTRS1984

Methane heat transfer investigation

Future high chamber pressure LOX/hydrocarbon booster engines require copper base alloy main combustion chamber coolant channels similar to the SSME to provide adequate cooling and reusable engine life. Therefore, it is of vital importance to evaluate the heat transfer characteristics and coking thresholds for LNG (94% methane) cooling, with a copper base alloy material adjacent to he fuel coolant. High pressure methane cooling and coking characteristics recently evaluated at Rocketdyne using stainless steel heated tubes at methane bulk temperatures and coolant wall temperatures typical of advanced engine operation except at lower heat fluxes as limited by the tube material. As expected, there was no coking observed. However, coking evaluations need be conducted with a copper base surface exposed to the methane coolant at higher heat fluxes approaching those of future high chamber pressure engines.

Source record↗
NASA NTRS1984

Methane heat transfer investigation

Future high chamber pressure LOX/hydrocarbon booster engines require copper-base alloy main combustion chamber coolant channels similar to the SSME to provide adequate cooling and resuable engine life. Therefore, it is of vital importance to evaluate the heat transfer characteristics and coking thresholds for LNG (94% methane) cooling, with a copper-base alloy material adjacent to the fuel coolant. High-pressure methane cooling and coking characteristics were recently evaluated using stainless-steel heated tubes at methane bulk temperatures and coolant wall temperatures typical of advanced engine operation except at lower heat fluxes as limited by the tube material. As expected, there was no coking observed. However, coking evaluations need be conducted with a copper-base surface exposed to the methane coolant at higher heat fluxes approaching those of future high chamber pressure engines.

Cook, R. T.↗
NASA NTRS1984

Heat transfer and turbulence measurements of a film-cooled flow over a convexly curved surface

Heat transfer and hydrodynamic boundary layer measurements were made in a turbulent boundary layer on a convex surface with 13 rows of injection. Both full- and partial-coverage cases were tested for three blowing ratios: 0.2, 0.4, and 0.6. Heat transfer results are discussed in terms of Stanton number with injection air temperature equal to wall temperature. In the blown region, for both m (blowing ratio) = 0.2 and 0.4, the local response of the boundary layer is dominated by curvature, and m = 0.4 shows higher cooling effects than m = 0.2 or 0.6, as expected, based on flat-plate results. Hydrodynamic measurements with a triple-hot-wire suggest the existence of two streamwise lanes. Lanes containing injection holes are highly affected by injection and show a definite streamwise evolution. Lanes which do not contain holes are not affected much by injection and show little change in the streamwise direction.

Furuhama, K.↗
NASA NTRS1990

A validation study of the Spark Navier Stokes code for nonreacting scramjet combustor flowfields

A detailed study of Mach 2 flow over a rearward facing step and staged normal injection behind a rearward facing step into a Mach 2 mainflow has been performed using the Spark three-dimensional Navier-Stokes code. Very good overall agreement was observed for the flowfield without normal injection. The notable exceptions were the velocity for which the measured values showed more scatter and the temperature at the wall for which the value measured was below the adiabatic wall temperature. Faithful representation of the incoming boundary layer was found to be important in achieving good agreement. For the staged jet case, the calculated and measured flow quantities showed reasonably good agreement, with two significant discrepancies: the calculated jet trajectory is inclined more to the vertical then the measured one, and the measured temperatures downstream of the jets are lower than the computed ones. Possible reasons for the discrepancies are suggested.

Eklund, Dean R.↗
NASA NTRS1950

Extension of boundary-layer heat-transfer theory to cooled turbine blades

An equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

HEAT TRANSFER, AERODYNAMIC↗
NASA NTRS2008

An Improved Formulation for Calorimetric Emittance Testing of Spacecraft Thermal Control Coatings

Spacecraft often really heavily on passive thermal control to maintain operating temperature. An important parameter in the spacecraft heat balance equation is the emittance of thermal control coatings as a function of coating temperature. One method for determining the emittance of spacecraft thermal control from elevated temperature to cryogenic temperatures relies on a calorimetric technique. The fundamental equation governing this test method can be found in numerous places in the literature and although it generally provides reasonable results, its formulation is based on a conceptual flaw that only becomes apparent when the sample temperature approaches the wall temperature during testing. This paper investigates the cause for this error and develops the correct formulation for calorimetric emittance testing. Experimental data will also be presented that illustrates the difference between the two formulations and the resulting difference in the calculated emittance.

Kauder, Lonny R.↗