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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Design Enhancements of the Two-Dimensional, Dual Throat Fluidic Thrust Vectoring Nozzle Concept

A Dual Throat Nozzle fluidic thrust vectoring technique that achieves higher thrust-vectoring efficiencies than other fluidic techniques, without sacrificing thrust efficiency has been developed at NASA Langley Research Center. The nozzle concept was designed with the aid of the structured-grid, Reynolds-averaged Navier-Stokes computational fluidic dynamics code PAB3D. This new concept combines the thrust efficiency of sonic-plane skewing with increased thrust-vectoring efficiencies obtained by maximizing pressure differentials in a separated cavity located downstream of the nozzle throat. By injecting secondary flow asymmetrically at the upstream minimum area, a new aerodynamic minimum area is formed downstream of the geometric minimum and the sonic line is skewed, thus vectoring the exhaust flow. The nozzle was tested in the NASA Langley Research Center Jet Exit Test Facility. Internal nozzle performance characteristics were defined for nozzle pressure ratios up to 10, with a range of secondary injection flow rates up to 10 percent of the primary flow rate. Most of the data included in this paper shows the effect of secondary injection rate at a nozzle pressure ratio of 4. The effects of modifying cavity divergence angle, convergence angle and cavity shape on internal nozzle performance were investigated, as were effects of injection geometry, hole or slot. In agreement with computationally predicted data, experimental data verified that decreasing cavity divergence angle had a negative impact and increasing cavity convergence angle had a positive impact on thrust vector angle and thrust efficiency. A curved cavity apex provided improved thrust ratios at some injection rates. However, overall nozzle performance suffered with no secondary injection. Injection holes were more efficient than the injection slot over the range of injection rates, but the slot generated larger thrust vector angles for injection rates less than 4 percent of the primary flow rate.

Flamm, Jeffrey D.↗
NASA NTRS1985

Semiempirical method of determining flow coefficients for pitot rake mass flow rate measurements

Flow coefficients applicable to area-weighted pitot rake mass flow rate measurements are presented for fully developed, turbulent flow in an annulus. A turbulent velocity profile is generated semiempirically for a given annulus hub-to-tip radius ratio and integrated numerically to determine the ideal mass flow rate. The calculated velocities at each probe location are then summed, and the flow rate as indicated by the rake is obtained. The flow coefficient to be used with the particular rake geometry is subsequently obtained by dividing the ideal flow rate by the rake-indicated flow rate. Flow coefficients ranged from 0.903 for one probe placed at a radius dividing two equal areas to 0.984 for a 10-probe area-weighted rake. Flow coefficients were not a strong function of annulus hub-to-tip radius ratio for rakes with three or more probes. The semiempirical method used to generate the turbulent velocity profiles is described in detail.

Trefny, C. J.↗
NASA NTRS1996

Turbulent flame propagation in partially premixed flames

Turbulent premixed flame propagation is essential in many practical devices. In the past, fundamental and modeling studies of propagating flames have generally focused on turbulent flame propagation in mixtures of homogeneous composition, i.e. a mixture where the fuel-oxidizer mass ratio, or equivalence ratio, is uniform. This situation corresponds to the ideal case of perfect premixing between fuel and oxidizer. In practical situations, however, deviations from this ideal case occur frequently. In stratified reciprocating engines, fuel injection and large-scale flow motions are fine-tuned to create a mean gradient of equivalence ratio in the combustion chamber which provides additional control on combustion performance. In aircraft engines, combustion occurs with fuel and secondary air injected at various locations resulting in a nonuniform equivalence ratio. In both examples, mean values of the equivalence ratio can exhibit strong spatial and temporal variations. These variations in mixture composition are particularly significant in engines that use direct fuel injection into the combustion chamber. In this case, the liquid fuel does not always completely vaporize and mix before combustion occurs, resulting in persistent rich and lean pockets into which the turbulent flame propagates. From a practical point of view, there are several basic and important issues regarding partially premixed combustion that need to be resolved. Two such issues are how reactant composition inhomogeneities affect the laminar and turbulent flame speeds, and how the burnt gas temperature varies as a function of these inhomogeneities. Knowledge of the flame speed is critical in optimizing combustion performance, and the minimization of pollutant emissions relies heavily on the temperature in the burnt gases. Another application of partially premixed combustion is found in the field of active control of turbulent combustion. One possible technique of active control consists of pulsating the fuel flow rate and thereby modulating the equivalence ratio (Bloxsidge et al. 1987). Models of partially premixed combustion would be extremely useful in addressing all these questions related to practical systems. Unfortunately, the lack of a fundamental understanding regarding partially premixed combustion has resulted in an absence of models which accurately capture the complex nature of these flames. Previous work on partially premixed combustion has focused primarily on laminar triple flames. Triple flames correspond to an extreme case where fuel and oxidizer are initially totally separated (Veynante et al. 1994 and Ruetsch et al. 1995). These flames have a nontrivial propagation speed and are believed to be a key element in the stabilization process of jet diffusion flames. Different theories have also been proposed in the literature to describe a turbulent flame propagating in a mixture with variable equivalence ratio (Muller et al. 1994), but few validations are available. The objective of the present study is to provide basic information on the effects of partial premixing in turbulent combustion. In the following, we use direct numerical simulations to study laminar and turbulent flame propagation with variable equivalence ratio.

Poinsot, T.↗
NASA NTRS2006

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a miss alignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery that was 0.01 lower than the experiment as was seen in the baseline. Comparison of the flow features the jet case revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Allan Brian G.↗
NASA NTRS2006

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

This paper will investigate the validation of the NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as a baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet experiment conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a free-stream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the fanface diameter. The numerical simulations with and without tunnel walls are performed, quantifying tunnel wall effects on the BLI inlet flow. A comparison is made between the numerical simulations and the BLI inlet experiment for the baseline and VG vane cases at various inlet mass flow rates. A comparison is also made to a BLI inlet jet configuration for varying actuator mass flow rates at a fixed inlet mass flow rate. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCP avg, very well within the designed operating range of the BLI inlet. A comparison of the average total pressure recovery showed that the simulations were able to predict the trends but had a negative 0.01 offset when compared to the experimental levels. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion levels for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a misalignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery offset that was 0.01 lower than the experiment as was seen in the baseline. Comparisons of the flow features for the jet cases revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Allan, Brian G.↗
NASA NTRS1973

Simultaneous film and convection cooling of a plate inserted in the exhaust stream of a gas turbine combustor

Data were obtained on a parallel-flow film- and convection-cooled test section placed in the exhaust stream of a rectangular-sector combustor. The combustor was operated at atmospheric pressure and at exhaust temperatures of 589 and 1033 K (600 and 1400 F). The cooling air was at ambient pressure and temperature. Test results indicate that it is better to use combined film and convection cooling rather than either film or convection cooling alone for a fixed total coolant flow. An optimum ratio of film to convection cooling flow rates was determined for the particular geometry tested. The experimental results compared well with calculated results.

Marek, C. J.↗
NASA NTRS1978

Pressure and heat-transfer distributions in a simulated wing-elevon cove with variable leakage at a free-stream Mach number of 6.9

An experimental aerodynamic heating investigation was conducted to determine effects of hot boundary-layer ingestion into the cove on the windward surface between a wing and elevon for cove seal leak areas nominally between 0 and 100 percent of cove entrance area. Pressure and heating-rate distributions were obtained on the wing and elevon surfaces and on the cove walls of a full-scale model that represented a section of the cove region on the space shuttle orbiter. Data were obtained for both attached and separated turbulent boundary layers upstream of the unswept cove entrance. Average free-stream Mach number was 6.9, average free-stream unit Reynolds numbers were 1.31 x 10 to the 6th power and 4.40 x 10 to the 6th power per meter (0.40 x 10 to the 6th power and 1.34 x 10 to the 6th power per foot), and average total temperature was 1888 K (3400 R). Cove pressures and heating rates varied as a function of seal leak area independent of leak aspect ratio. Although cove heating rates for attached flow did not appear intolerable, it was postulated that convective heating in the cove may increase with time. For separated flow, the cove environment was considered too severe for unprotected interior structures of control surfaces.

Deveikis, W. D.↗
NASA NTRS1986

Evaluation of a hybrid hydrostatic bearing for cryogenic turbopump application

A hybrid hydrostatic bearing was designed to operate in liquid hydrogen at speeds to 80,000 rpm and radial loads to 440 n (100 lbf). The bearing assembly consisted of a pair of 20-mm angular-contact ball bearings encased in a journal, which was in turn supported by a fluid film of liquid hydrogen. The size and operating conditions of the bearing were selected to be compatible with the operating requirements of an advanced technology turbopump. Several test parameters were varied to characterize the bearing's steady-state operation. The rotation of the tester shaft was varied between 0 and 80,000 rpm. Bearing inlet fluid pressure was varied between 2.07 and 4.48 MPa (300 and 650 psia), while the fluid sump pressure was independently varied between 0.34 and 2.07 MPa (50 and 300 psia). The maximum radial load applied to the bearing was 440 N (110 lbf). Measured hybrid-hydrostatic-bearing stiffness was 1.5 times greater than predicted, while the fluid flow rate through the bearing was 35 to 65 percent less than predicted. Under two-phase fluid conditions, the stiffness was even greater and the flow rate was less. The optimal pressure ratio for the bearing should be between 0.2 and 0.55 depending on the balance desired between bearing efficiency and stiffness. Startup and shutdown cyclic tests were conducted to demonstrate the ability of the hybrid-hydrostatic-bearing assembly to survive at least a 300-firing-duty cycle. For a typical cycle, the shaft was accelerated to 50,000 rpm in 1.8 sec. The bearing operated for 337 start-stop cycles without failure.

Spica, P. W.↗
NASA NTRS1987

Amorphous silicon deposition diagnostics using coherent anti-Stokes Raman spectroscopy

This paper reports on an in situ silicon deposition process diagnostics, using CARS performed at state-of-the-art a-Si:H film deposition conditions in a reactor designed for a-Si:H solar cell fabrication. The diagnostics procedure measures the silane depletion in an RF plasma as a function of the silane flow rate (where the relationship is linear between 50-percent depletion, at a flow rate of 5.6 sccm, and about 8-percent depletion, at 80 sccm) and the RF power (where the silane depletion is linearly dependent on the RF power in the region of 4 to 12 W). The linear RF power dependence of the silane depletion is considered to be consistent with the mechanism of silane decomposition primarily by electron impact dissociations, while the flow rate dependence is interpreted in terms of the residence time of the SiH4 molecules in the glow discharge region, where an increase of the residence time at a low flow rate results in a high depletion ratio.

Shing, Y. H.↗
NASA NTRS1992

The effect of inlet swirl on the rotordynamic shroud forces in a centrifugal pump

The role played by fluid forces in determining the rotordynamic stability of a centrifugal pump is gaining increasing attention. The present research investigates the contributions to the rotordynamic forces from the discharge-to-suction leakage flows between the front shroud of the rotating impeller and the stationary pump casing. In particular, the dependency of the rotordynamic characteristics of leakage flows on the swirl at the inlet to the leakage path was examined. An inlet guide vane was designed for the experiment so that swirl could be introduced at the leakage flow inlet. The data demonstrates substantial rotordynamic effects and a destabilizing tangential force for small positive whirl ratios; this force decreased with increasing flow rate. The effect of swirl on the rotordynamic forces was found to be destabilizing.

Ginzburg, A.↗
NASA NTRS1993

A viscous shock-layer analysis of 2-D and axisymmetric flows

Results are obtained for cylindrical leading edges of proposed transatmospheric vehicles by employing a 2D viscous shock-layer code for nonequilibrium and perfect gas flows. The accuracy and efficiency of the planar code is verified through detailed comparisons with other predictions. It is found to be as accurate and robust as its axisymmetric counterpart. This study includes results for nose radii ranging from 0.01 to 2.0 ft and half-angles of 5 and 6 deg for both cylindrically-blunted wedges and spherically blunted cones (included for comparison). Some results are presented as a ratio of the noncatalytic to the corresponding fully catalytic heating value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flows. This ratio and the individual heating rates are smaller for cylindrically-blunted wedges with small nose radii, relative to the spherically-blunted cones (at the same radius). Therefore, a larger potential exists for heating reduction in cylindrically-blunted as compared with the spherically-blunted surfaces for finite-rate chemistry.

Gupta, Roop N.↗
NASA NTRS1952

Short-bearing approximation for full journal bearings

A short-bearing approximation of pressure distribution in the oil film is presented which is an extension of the pressure-distribution function of Michell and Cardullo and includes end-leakage effects. Equations giving applied load, attitude angle, location and magnitude of peak film pressure, friction, and required oil flow rate as functions of the eccentricity ratio are also given. The capacity number, a basic non dimensional quantity resulting from this analysis is the product of the Sommerfeld number and the square of the length-diameter ratio. Curves determined by this analysis are compared with previously published experimental data and theoretical curves of Sommerfeld and Cameron and Wood. Conclusions reached indicate that this approximation is of practical value for analysis of short bearings.

FRICTION AND LUBRICATION - HYDRODYNAMIC THEORY↗
NASA NTRS1995

Low-state disks and low-beta disks

Stellar black hole candidates (BHCs) exhibit bimodal spectral states. We calculate nonthermal disk spectra, demonstrating that a large photon index (alpha (sub x) approximately 2-3) observed in the soft (high) state is due to a copious soft photon supply, whereas soft photon starvation leads to a smaller index (alpha (sub x) approximately 1.5-2) in the hard (low) state. Thus, the absence of the soft component flux in the low state cannot be due to obscuration. A possible disk configuration during the low state is discussed. We proposed that a low-state disk may be a low-beta disk in which magnetic pressure may exceed gas pressure becuase of the suppression of field escape by a strong shear. As a result, disk material will take the form of blobs constricted by mainly toroidal magnetic fields. Fields are dissipated mainly by occasional reconnection events with a huge energy release. This will account for large-amplitude, aperiodic X-ray variations (flickering) and high-energy radiation with small alpha(sub x) from hard state BHCs and possibly from active galactic nuclei. Further, we propose a hysteretic relation between the mass-flow rate and plasma-beta, a ratio of gas pressure to magnetic pressure, for the spectral evolution of transient BHCs. The disk is in the low-beta state in quiescence and early rise. The low-beta disk is optically thin and affected by advection. A hard-to-soft transition occurs before the peak luminosity, since there is no advection-dominated branch at higher luminosities. An optically thick, high-beta disk appears at small radii. In the decay phase of the light curve, the standard-type disk becomes effectively optically thin, when a soft-hard transition is triggered. High-beta plasmas in the main body shrink to form minute blobs, and low-beta coronal plasma fills interblob space.

Mineshige, Shin↗
NASA NTRS2013

Component Testing of the J-2X Augmented Spark Igniter (ASI)

In support of the development of the J-2X engine, 201 low pressure, liquid oxygen / liquid hydrogen (LOX/LH2) J-2X Augmented Spark Igniter (ASI) subsystem ignition tests were conducted at Marshall Space Flight Center (MSFC). The main objective of these tests was to start the ASI within the anticipated J-2X engine start box, as well as outside of it, to check for ignition margin. The setup for the J-2X ASI component testing simulated, as much as possible, the tank-head start-up configuration of the ASI within the J-2X Engine. The ignition tests were divided into 124 vacuum start tests to simulate altitude start on a flight engine, and 77 sea-level start tests to simulate the first set of ground tests for the J-2X Engine at Stennis Space Center (SSC). Other ignition parameters that were varied included propellant tank pressures, oxidizer temperature entering the ASI oxidizer feedline, oxidizer valve timing, spark igniter condition (new versus damaged), and oxidizer and fuel feedline orifice sizes. Propellant blowdowns using venturis sized to simulate the ASI resistance allowed calculation of transient propellant mass flow rates as well as global mixture ratio for all ignition tests. Global mixture ratio within the ASI at the time of ignition varied from 0.2 to 1.2. Detailed electronics data obtained from an instrumented ignition lead allowed characterization of the breakdown voltage, sustaining voltage and energy contained in each spark as the ASI propellants ignited. Results indicated that ignition always occurred within the first five sparks when both propellants were present in the ASI chamber.

Osborne, Robin J.↗
NASA NTRS2017

Progress of High Efficiency Centrifugal Compressor Simulations Using TURBO

Three-dimensional, time-accurate, and phase-lagged computational fluid dynamics (CFD) simulations of the High Efficiency Centrifugal Compressor (HECC) stage were generated using the TURBO solver. Changes to the TURBO Parallel Version 4 source code were made in order to properly model the no-slip boundary condition along the spinning hub region for centrifugal impellers. A startup procedure was developed to generate a converged flow field in TURBO. This procedure initialized computations on a coarsened mesh generated by the Turbomachinery Gridding System (TGS) and relied on a method of systematically increasing wheel speed and backpressure. Baseline design-speed TURBO results generally overpredicted total pressure ratio, adiabatic efficiency, and the choking flow rate of the HECC stage as compared with the design-intent CFD results of Code Leo. Including diffuser fillet geometry in the TURBO computation resulted in a 0.6 percent reduction in the choking flow rate and led to a better match with design-intent CFD. Diffuser fillets reduced annulus cross-sectional area but also reduced corner separation, and thus blockage, in the diffuser passage. It was found that the TURBO computations are somewhat insensitive to inlet total pressure changing from the TURBO default inlet pressure of 14.7 pounds per square inch (101.35 kilopascals) down to 11.0 pounds per square inch (75.83 kilopascals), the inlet pressure of the component test. Off-design tip clearance was modeled in TURBO in two computations: one in which the blade tip geometry was trimmed by 12 mils (0.3048 millimeters), and another in which the hub flow path was moved to reflect a 12-mil axial shift in the impeller hub, creating a step at the hub. The one-dimensional results of these two computations indicate non-negligible differences between the two modeling approaches.

turbomachinery↗
NASA NTRS1988

Leakage predictions for Rayleigh-step, helium-purge seals

Rayleigh-step, helium purge, annular shaft seals, studied for use in liquid oxygen turbopumps, generate a hydrodynamic force that enables the seal to follow shaft perturbations. Hence, smaller clearances can be used to reduce seal leakage. FLOWCAL, a computer code developed by Mechanical Technology Incorporated, predicts gas flow rate through an annular seal with an axial pressure gradient. Analysis of a 50-mm Rayleigh-step, helium-purge, annular seal showed the flow rate increased axial pressure gradient, downstream pressure, and eccentricity ratio. Increased inlet temperature reduced leakage. Predictions made at maximum and minimum clearances (due to centrifugal and thermal growths, machining tolerances and + or - 2 percent uncertainty in the clearance measurement) placed wide boundaries on expected flow rates. The widest boundaries were set by thermal growth conditions. Predicted flow rates for a 50-mm Rayleigh-step, helium-purge, annular seal underestimated measured flow rates by three to seven times. However, the analysis did accurately predict flow rates for choked gas flow through annular seals when compared to flow rates measured in two other independent studies.

Proctor, Margaret P.↗
NASA NTRS1995

Hybrid fuel formulation and technology development

The objective was to develop an improved hybrid fuel with higher regression rate, a regression rate expression exponent close to 0.5, lower cost, and higher density. The approach was to formulate candidate fuels based on promising concepts, perform thermomechanical analyses to select the most promising candidates, develop laboratory processes to fabricate fuel grains as needed, fabricate fuel grains and test in a small lab-scale motor, select the best candidate, and then scale up and validate performance in a 2500 lbf scale, 11-inch diameter motor. The characteristics of a high performance fuel have been verified in 11-inch motor testing. The advanced fuel exhibits a 15% increase in density over an all hydrocarbon formulation accompanied by a 50% increase in regression rate, which when multiplied by the increase in density yields a 70% increase in fuel mass flow rate; has a significantly lower oxidizer-to-fuel (O/F) ratio requirement at 1.5; has a significantly decreased axial regression rate variation making for more uniform propellant flow throughout motor operation; is very clean burning; extinguishes cleanly and quickly; and burns with a high combustion efficiency.

Dean, D. L.↗
NASA NTRS1998

PAB3D Simulations of a Nozzle with Fluidic Injection for Yaw Thrust-Vector Control

An experimental and computational study was conducted on an exhaust nozzle with fluidic injection for yaw thrust-vector control. The nozzle concept was tested experimentally in the NASA Langley Jet Exit Test Facility (JETF) at nozzle pressure ratios up to 4 and secondary fluidic injection flow rates up to 15 percent of the primary flow rate. Although many injection-port geometries and two nozzle planforms (symmetric and asymmetric) were tested experimentally, this paper focuses on the computational results of the more successful asymmetric planform with a slot injection port. This nozzle concept was simulated with the Navier-Stokes flow solver, PAB3D, invoking the Shih, Zhu, and Lumley algebraic Reynolds stress turbulence model (ASM) at nozzle pressure ratios (NPRs) of 2,3, and 4 with secondary to primary injection flow rates (w(sub s)/w(sub p)) of 0, 2, 7 and 10 percent.

Deere, Karen A.↗