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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 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
NASA NTRS1943

Heat Transfer Over the Circumference of a Heated Cylinder in Transverse Flow

A method for recording the local heat-transfer coefficients on bodies in flow was developed. The cylinder surface was kept at constant temperature by the condensation of vapor except for a narrow strip which is heated separately to the same temperature by electricity. The heat-transfer coefficient at each point was determined from the electric heat output and the temperature increase. The distribution of the heat transfer along the circumference of cylinders was recorded over a range of Reynolds numbers of from 5000 to 426,000. The pressure distribution was measured at the same time. At Reynolds numbers up to around 100,000 high maximums of the heat transfer occurred in the forward stagnation point at and on the rear side at 180C, while at around 80 the heat-transfer coefficient on both sides of the cylinder behind the forward stagnation point manifested distinct minimums. Two other maximums occurred at around 115 C behind the forward stagnation point between 170,000 and 426,000. At 426,000 the heat transfer at the location of those maximums was almost twice as great as in the forward stagnation point, and the rear half of the cylinder diffused about 60 percent of the entire heat, The tests are compared with the results of other experimental and theoretical investigations.

Schmidt, Ernst↗
NASA NTRS1991

Transonic Navier-Stokes Solutions about a Complex High-Speed Accelerator Configuration

Three dimensional transonic viscous flow computations are presented for a generic high-speed accelerator model which includes wing, body, filets,and a no-flow through engine nacelle. solutions are obtained from an algorithm for the compressible Navier-Stokes equations which incorporated an upwind-biased, flux-vector-splitting approach along with longitudinally-patched grids. Results are presented for fully turbulent flow assumptions and include correlations with wind tunnel data. A good quantitative agreement for the forebody surface pressure distribution is achieved between computations and the available wind-tunnel measurements at M(sub infinity) = 0.9. Furthermore, it is demonstrated that the flow is stagnating around the boattail region due to separation from the aft-engine cowl lip.

Ghaffari, F.↗
NASA NTRS2012

Nonequilibrium Ablation of Phenolic Impregnated Carbon Ablator

In previous work, an equilibrium ablation and thermal response model for Phenolic Impregnated Carbon Ablator was developed. In general, over a wide range of test conditions, model predictions compared well with arcjet data for surface recession, surface temperature, in-depth temperature at multiple thermocouples, and char depth. In this work, additional arcjet tests were conducted at stagnation conditions down to 40 W/sq cm and 1.6 kPa. The new data suggest that nonequilibrium effects become important for ablation predictions at heat flux or pressure below about 80 W/sq cm or 10 kPa, respectively. Modifications to the ablation model to account for nonequilibrium effects are investigated. Predictions of the equilibrium and nonequilibrium models are compared with the arcjet data.

Milos, Frank S.↗
NASA NTRS1972

An experimental investigation of turbulent boundary layers at high Mach number and Reynolds numbers

Skin friction, heat transfer and pressure measurements were obtained in laminar, transitional and turbulent boundary layers on flat plates at Mach numbers from 7 to 13 at wall-to-free stream stagnation temperature ratios from 0.1 to 0.3. Measurements in laminar flows were in excellent agreement with the theory of Cheng. Correlations of the transition measurements with measurements on flight vehicles and in ballistic ranges show good agreement. Our transition measurements do not correlate well with those of Pate and Schueler. Comparisons have been made between the skin friction and heat transfer measurements and the theories of Van Driest, Eckert and Spalding and Chi. These comparisons reveal in general that at the high end of our Mach number range (10-13) the theory of Van Driest is in best agreement with the data, whereas at lower Mach numbers (6.5-10) the Spalding Chi theory is in better agreement with the measurements.

Holden, M. S.↗
NASA NTRS1974

Use of potential flow theory to evaluate subsonic inlet data from a simulator-powered nacelle at cruise conditions

Incompressible potential flow theory corrected for compressibility effects, using the Lieblein-Stockman compressibility correction, was used to predict surface and flow field static pressures for a subsonic inlet at cruise conditions. The calculated internal and external surface static pressures were in good agreement with data at most conditions. The analysis was used to determine the capture stream-tube location and static-pressure distribution. Additive drag coefficients obtained from these results were consistently higher than those obtained using one-dimensional compressible flow theory. Increasing the distance between the inlet and boattail increased the cowl drag force. The effect of the boundary layer on internal and external surface static-pressure distributions was small at the design cruise condition. The analytical results may be used as an aid to data reduction and for predicting inlet mass flow, stagnation point location, and inlet additive drag.

Bober, L. J.↗
NASA NTRS1990

A unique high heat flux facility for testing hypersonic engine components

This paper describes the Hot Gas Facility, a unique, reliable, and cost-effective high-heat-flux facility for testing hypersonic engine components developed at the NASA Lewis Research Center. The Hot Gas Facility is capable of providing heat fluxes ranging from 200 Btu/sq ft per sec on flat surfaces up to 8000 Btu/sq ft per sec at a leading edge stagnation point. The usefulness of the Hot Gas Facility for the NASP community was demonstrated by testing hydrogen-cooled structures over a range of temperatures and pressures. Ranges of the Reynolds numbers, Prandtl numbers, enthalpy, and heat fluxes similar to those expected during hypersonic flights were achieved.

Melis, Matthew E.↗
NASA NTRS2002

Dual-Mode Combustion

The Department of Mechanical and Aerospace Engineering at the University of Virginia has conducted an investigation of the mixing and combustion processes in a hydrogen fueled dual-mode scramjet combustor. The experiment essentially consisted of the "direct connect" continuous operation of a Mach 2 rectangular combustor with a single unswept ramp fuel injector. The stagnation enthalpy of the test flow simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and laser based diagnostics. These diagnostics included, pressure and wall temperature measurements, Fuel Plume Imaging (FPI) and Particle Image Velocimetry (PIV). A schematic of the combustor configuration and a summary of the measurements obtained are presented. The experimental work at UVa was parallel by Computational Fluid Dynamics (CFD) work at NASA Langley. The numerical and experiment results are compared in this document.

Goyne, Christopher P.↗
DOE OSTI2024

Formation of a spherical plasma liner for plasma-jet-driven magneto-inertial fusion

Plasma-jet-driven magneto-inertial fusion is an alternative approach to controlled nuclear fusion, which aims to utilize a line-replaceable dense plasma liner as a repetitive spherical compression driver. In this experiment, first measurements of the formation of a spherical argon plasma liner formed from 36 discrete pulsed plasma jets are obtained on the Plasma Liner Experiment. Properties including liner uniformity and morphology, plasma density, temperature, and ram pressure are assessed as a function of time throughout the implosion process and indicate an apparent transition from initial kinetic inter-jet interpenetration to collisional regime near stagnation times, in accordance with theoretical expectation. A lack of primary shock structures between adjacent jets during flight implies that arbitrarily smooth liners may be formed by way of corresponding improvements in jet parameters and control. The measurements facilitate the benchmarking of computational models and understanding the scaling of plasma liners toward fusion-relevant energy density.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1985

Measurements in the flow field of a cylinder with a laser transit anemometer and a drag rake in the Langley 0.3 m transonic cryogenic tunnel

An experiment was conducted at the 0.3 m Transonic Cryogenic Tunnel using a Laser Transit Anemometer (LTA) to probe the flow field around a 3.05 centimeter-diameter circular cylinder. Measurements were made along the stagnation line and determination of particle size was evaluated by their ability to follow the flow field. The LTA system was also used to scan into the boundary layer near the 45 degree point on the model. Results of these scans are presented in graphic and tabular form. Flow field particle seeding was accomplished by inbleeding dry service air. The residual moisture (5-10 ppm) condensed and formed ice particles which served as Mie scattering centers for the LTA detection system. Comparison of data taken along the stagnation streamline with theory indicated that these particles tracked the velocity gradient of the flow. Tunnel operating conditions for the tests were a Mach number of 0.3, a pressure of 1.93 x 100000 n/m squared, and a temperature of 225 degrees K. Free stream Mach number and pressure were varied for the particle size determination.

Honaker, W. C.↗
NASA NTRS1990

Transonic Navier-Stokes solutions about a complex high-speed accelerator configuration

Three-dimensional transonic viscous flow computations are presented for a generic high-speed accelerator model which includes wing, body, fillets, and a no-flow through engine nacelle. Solutions are obtained from an algorithm for the compressible Navier-Stokes equations which incorporates an upwind-biased, flux-vector-splitting approach along with longitudinally-patched grids. Results are presented for fully turbulent flow assumptions and include correlations with wind tunnel data. A good quantitative agreement for the forebody surface pressure distribution is achieved between computations and the available wind-tunnel measurements at freestream Mach number equal to 0.9. Furthermore, it is demonstrated that the flow is stagnating around the boattail region due to separation from the aft-engine cowl lip.

Ghaffari, Farhad↗
NASA NTRS1994

Analysis and design of numerical schemes for gas dynamics. 2: Artificial diffusion and discrete shock structure

The effect of artificial diffusion on discrete shock structures is examined for a family of schemes which includes scalar diffusion, convective upwind and split pressure (CUSP) schemes, and upwind schemes with characteristics splitting. The analysis leads to conditions on the diffusive flux such that stationary discrete shocks can contain a single interior point. The simplest formulation which meets these conditions is a CUSP scheme in which the coefficients of the pressure differences is fully determined by the coefficient of convective diffusion. It is also shown how both the characteristic and CUSP schemes can be modified to preserve constant stagnation enthalpy in steady flow, leading to four variants, the E and H-characteristic schemes, and the E and H-CUSP schemes. Numerical results are presented which confirm the properties of these schemes.

Jameson, Antony↗
NASA NTRS2019

Physical Simulation of Rocket Exhaust Aerodynamics Using Heated Ethane: Conceptual Foundations

Heated ethane (C2H6) has been proposed as an alternative to inert gases for use as a motive fluid in the experimental simulation of rocket exhaust plumes. By adjusting stagnation temperature, the isentropic exponent of ethane can be tuned to approximate those produced by common rocket propellants including hydrogen, hypergols, alcohols, and hydrocarbons. As a result, ethane can be made to follow a nozzle expansion process which is nearly identical to realistic rocket engine flow fields. Additionally, its high auto-ignition temperature and resistance to condensation enable the testing of expansion ratios much larger than conventional inertgas testing. NASA SSC has performed quasi-one-dimensional analyses using the Chemical Equilibrium with Applications (CEA) code as a preliminary means to compare flow fields produced by non-reacting ethane to those of reacting combustion products. A LO2/LH2 rocket engine operating at a chamber pressure of 5.0 MPa and a mixture ratio of 6.1 was used as an example case to demonstrate ethane’s efficacy as a simulant. Errors for key similarity parameters were compared to legacy cold-flow test methods. Additional errors induced by machining tolerances and chemical impurities were also examined. Results suggest that at a 3% geometric scale and ~500 K ethane stagnation temperature, an error of less than 2.5% throughout the flow field is realistically achievable along the dimensions of Mach number, Reynolds number, pressure ratio, and isentropic exponent. The development of an experimental test bed for validation of this configuration is currently underway.

Jones, Daniel R.↗
NASA NTRS2011

Numerical Prediction of Radiation Measurements Taken in the X2 Facility for Mars and Titan Gas Mixtures

Thermochemical relaxation behind a normal shock in Mars and Titan gas mixtures is simulated using a CFD solver, DPLR, for a hemisphere of 1 m radius; the thermochemical relaxation along the stagnation streamline is considered equivalent to the flow behind a normal shock. Flow simulations are performed for a Titan gas mixture (98% N2, 2% CH4 by volume) for shock speeds of 5.7 and 7.6 km/s and pressures ranging from 20 to 1000 Pa, and a Mars gas mixture (96% CO2, and 4% N2 by volume) for a shock speed of 8.6 km/s and freestream pressure of 13 Pa. For each case, the temperatures and number densities of chemical species obtained from the CFD flow predictions are used as an input to a line-by-line radiation code, NEQAIR. The NEQAIR code is then used to compute the spatial distribution of volumetric radiance starting from the shock front to the point where thermochemical equilibrium is nominally established. Computations of volumetric spectral radiance assume Boltzmann distributions over radiatively linked electronic states of atoms and molecules. The results of these simulations are compared against experimental data acquired in the X2 facility at the University of Queensland, Australia. The experimental measurements were taken over a spectral range of 310-450 nm where the dominant contributor to radiation is the CN violet band system. In almost all cases, the present approach of computing the spatial variation of post-shock volumetric radiance by applying NEQAIR along a stagnation line computed using a high-fidelity flow solver with good spatial resolution of the relaxation zone is shown to replicate trends in measured relaxation of radiance for both Mars and Titan gas mixtures.

Palmer, Grant↗
DOE OSTI2021

Three dimensional low-mode areal-density non-uniformities in indirect-drive implosions at the National Ignition Facility

To achieve hotspot ignition, an inertial confinement fusion implosion must achieve high hotspot pressure that is inertially confined by a dense shell of DT fuel. This requires a symmetric implosion having high in-flight shell velocity and high areal density at stagnation. The size of the driver and scale of the capsule required can be minimized by maintaining a high efficiency of energy coupling from the imploding shell to the hotspot. Significant 3D low mode asymmetries, however, are commonly observed in indirect-drive implosions and reduce the coupling of shell kinetic energy to the hotspot. To better quantify the magnitudes and impacts of shell density asymmetries, we have developed new analysis techniques and analytic models [Hurricane et al., Phys. Plasmas 27(6), 062704 (2020)]. To build confidence in the underlying data, we have also developed an analytic neutron transport model to cross-compare two independent measurements of asymmetry, which shows excellent agreement across shots for mode-1 (ℓ = 1). This work also demonstrates that asymmetry can introduce potential sampling bias into down-scattered ratio measurements causing the solid-angle-average and uncertainty-weighted-average down-scattered ratios to differ significantly. Diagnosing asymmetries beyond mode-1 (ℓ > 1) presents significant challenges. Using new diagnostic instruments and analysis techniques, however, evidence of significant Legendre mode P2 (ℓ = 2, m = 0) and additional 3D asymmetries (ℓ > 1, m ≠ 0) are beginning to emerge from the high precision activation diagnostic data (real-time nuclear activation detectors) and down-scattered neutron imaging data.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1982

Numerical modeling and optimization of polymer melt processing operations

The application of finite element computer analyses to polymer flows of the type encountered in melt processing operations is described. A code capable of predicting values of fluid veleocity, pressure, shear stress, and temperature at any point within the flow field was developed. As such, is is of value in diagnosing such processing problems as regions of fluid stagnation at which thermal degradation may occur, or regions of excessive shear deformation which lead to thermomechanical damage. It is further able to generate predictions of the forces which must be applied to the melt to achieve the desired flow, and this information is of value in designing processing equipment of optimal efficiency and minimum energy consumption.

Roylance, D.↗
NASA NTRS1988

Thermal analysis of thermal barrier coatings in a high heat flux environment

Gas temperatures and pressures were measured around the second test position in the H2/O2 rocket engine at NASA-Lewis. Measured gas temperatures generally varied from 1210 to 1390 C. Measured pressures were in good agreement with other studies for throat tubes in a square chamber rocket engine. Heat transfer coefficients were measured at 90 and 180 degrees from the stagnation point and resulted in values of 27.5 and 8.5 kW/sq m C, respectively. A thermal model was developed to predict temperatures in bare and coated tubes and rods. Agreement between measured and predicted temperatures below the surface of a bare Mar-M 246 tube was very good for most of the heat up and cool down period. Predicted temperatures were significantly below measured temperatures for the coated tubes. A thermal model to simulate heat transfer to the leading edge of an HPFTP blade was developed and showed that TBCs can significantly dampen the thermal transient which occurs in the HPFTP during the startup of the SSME.

Nesbitt, James A.↗
NASA NTRS1959

Measurements of Local Heat Transfer and Pressure on Six 2-Inch-Diameter Blunt Bodies at a Mach Number of 4.95 and at Reynolds Numbers Per Foot up to 81 x 10(exp 6)

Measurements of the local heat transfer and pressure distribution have been made on six 2-inch-diameter, blunt, axially symmetric bodies in the Langley gas dynamics laboratory at a Mach number of 4.95 and at Reynolds numbers per foot up to 81 x 10(exp 6). During the investigation laminar flow was observed over a hemispherical-nosed body having a surface finish from 10 to 20 microinches at the highest test Reynolds number per foot (for this configuration) of 77.4 x 10(exp 6). Though it was repeatedly possible to measure completely laminar flow at this Reynolds number for the hemisphere, it was not possible to observe completely laminar flow on the flat-nosed body for similar conditions. The significance of this phenomenon is obscured by the observation that the effects of particle impacts on the surface in causing roughness were more pronounced on the flat-nosed body. For engineering purposes, a method developed by M. Richard Dennison while employed by Lockheed Aircraft Corporation appears to be a reasonable procedure for estimating turbulent heat transfer provided transition occurs at a forward location on the body. For rearward-transition locations, the method is much poorer for the hemispherical nose than for the flat nose. The pressures measured on the hemisphere agreed very well with those of the modified Newtonian theory, whereas the pressures on all other bodies, except on the flat-nosed body, were bracketed by modified Newtonian theory both with and without centrifugal forces. For the hemisphere, the stagnation-point velocity gradient agreed very well with Newtonian theory. The stagnation-point velocity gradient for the flat- nosed model was 0.31 of the value for the hemispherical-nosed model. If a Newtonian type of flow is assumed, the ratio 0.31 will be independent of Much number and real-gas effects.

Cooper, Morton↗
NASA NTRS1979

A careful numerical study of flowfields about external conical corners. I - Symmetric configurations

A numerical study of the flowfield about symmetrical external axial corners formed by the juncture of swept compressive wedges is presented. The geometrical configuration under investigation allows a unified treatment of external corners typical of delta wings and of rectangular inlets. Comparisons are made with other numerical results. For the delta wing configuration, the occurrence of an anomalous shock behavior predicted by Gonor is discussed. A parametric study shows the singular behavior of the pressure at the corner as a function of the external corner angle, and a second parametric study shows the effect of finite corner radius on the location of the crossflow stagnation point. Previous theoretical predictions of certain flow features typical of corner flows agree well with present numerical results.

Salas, M. D.↗