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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Viscous analyses for flow through subsonic and supersonic intakes

A parabolized Navier-Stokes code was used to analyze a number of diffusers typical of a modern inlet design. The effect of curvature of the diffuser centerline and transitioning cross sections was evaluated to determine the primary cause of the flow distortion in the duct. Results are presented for S-shaped intakes with circular and transitioning cross sections. Special emphasis is placed on verification of the analysis to accurately predict distorted flow fields resulting from pressure-driven secondary flows. The effect of vortex generators on reducing the distortion of intakes is presented. Comparisons of the experimental and analytical total pressure contours at the exit of the intake exhibit good agreement. In the case of supersonic inlets, computations of the inlet flow field reveal that large secondary flow regions may be generated just inside of the intake. These strong flows may lead to separated flow regions and cause pronounced distortions upstream of the compressor.

Povinelli, Louis A.↗
NASA NTRS2000

Global Design Optimization for Fluid Machinery Applications

Recent experiences in utilizing the global optimization methodology, based on polynomial and neural network techniques for fluid machinery design are summarized. Global optimization methods can utilize the information collected from various sources and by different tools. These methods offer multi-criterion optimization, handle the existence of multiple design points and trade-offs via insight into the entire design space can easily perform tasks in parallel, and are often effective in filtering the noise intrinsic to numerical and experimental data. Another advantage is that these methods do not need to calculate the sensitivity of each design variable locally. However, a successful application of the global optimization method needs to address issues related to data requirements with an increase in the number of design variables and methods for predicting the model performance. Examples of applications selected from rocket propulsion components including a supersonic turbine and an injector element and a turbulent flow diffuser are used to illustrate the usefulness of the global optimization method.

Shyy, Wei↗
NASA NTRS2011

Generalized Jeans' Escape of Pick-Up Ions in Quasi-Linear Relaxation

Jeans escape is a well-validated formulation of upper atmospheric escape that we have generalized to estimate plasma escape from ionospheres. It involves the computation of the parts of particle velocity space that are unbound by the gravitational potential at the exobase, followed by a calculation of the flux carried by such unbound particles as they escape from the potential well. To generalize this approach for ions, we superposed an electrostatic ambipolar potential and a centrifugal potential, for motions across and along a divergent magnetic field. We then considered how the presence of superthermal electrons, produced by precipitating auroral primary electrons, controls the ambipolar potential. We also showed that the centrifugal potential plays a small role in controlling the mass escape flux from the terrestrial ionosphere. We then applied the transverse ion velocity distribution produced when ions, picked up by supersonic (i.e., auroral) ionospheric convection, relax via quasi-linear diffusion, as estimated for cometary comas [1]. The results provide a theoretical basis for observed ion escape response to electromagnetic and kinetic energy sources. They also suggest that super-sonic but sub-Alfvenic flow, with ion pick-up, is a unique and important regime of ion-neutral coupling, in which plasma wave-particle interactions are driven by ion-neutral collisions at densities for which the collision frequency falls near or below the gyro-frequency. As another possible illustration of this process, the heliopause ribbon discovered by the IBEX mission involves interactions between the solar wind ions and the interstellar neutral gas, in a regime that may be analogous [2].

Moore, T. E.↗
NASA NTRS1987

Combustion rate limits of hydrogen plus hydrocarbon fuel: Air diffusion flames from an opposed jet burner technique

Combustion of H2/hydrocarbon (HC) fuel mixtures may be considered in certain volume-limited supersonic airbreathing propulsion applications. Effects of HC addition to H2 were evaluated, using a recent argon-bathed, coaxial, tubular opposed jet burner (OJB) technique to measure the extinction limits of counterflow diffusion flames. The OJB flames were formed by a laminar jet of (N2 and/or HC)-diluted H2 mixture opposed by a similar jet of air at ambient conditions. The OJB data, derived from respective binary mixtures of H2 and methane, ethylene, or propane HCs, were used to characterize BLOWOFF and RESTORE. BLOWOFF is a sudden breaking of the dish-shaped OJB flame to a stable torus or ring shape, and RESTORE marks sudden restoration of the central flame by radial inward flame propagation. BLOWOFF is a measure of kinetically-limited flame reactivity/speed under highly stretched, but relatively ideal impingement flow conditions. RESTORE measures inward radial flame propagation rate, which is sensitive to ignition processes in the cool central core. It is concluded that relatively small molar amounts of added HC greatly reduce the reactivity characteristics of counterflow hydrogen-air diffusion flames, for ambient initial conditions.

Pellett, Gerald L.↗
NASA NTRS1981

Advanced supersonic transport propulsion and configuration technology improvements

This paper presents the results of recent Douglas Aircraft Company integration studies for an advanced supersonic transport. The studies include technology improvements such as superplastic formed and diffusion-bonded titanium sandwich primary structure, composite secondary structure, bicone engine inlet, improved mechanical suppressors based on flight test data, improved aerodynamic efficiency based on wind tunnel test data, and updated performance for both variable-cycle and low-bypass-ratio engines. Technology development requirements for an economically viable and environmentally acceptable advanced supersonic transport are defined through these studies. A new Douglas Aircraft Company baseline supersonic transport designed to carry 350 passengers is defined for cost comparisons.

Rowe, W. T.↗
DOE OSTI2024

A Variable Eddington Factor Model for Thermal Radiative Transfer with Closure Based on Data-Driven Shape Function

Here, a new variable Eddington factor (VEF) model is presented for nonlinear problems of thermal radiative transfer (TRT). The VEF model is data-driven and acts on known (a-priori) radiation-diffusion solutions for material temperatures in the TRT problem. A linear auxiliary problem is constructed for the radiative transfer equation (RTE) whose emission source and opacities are evaluated at these known material temperatures. The solution to this RTE approximates the specific intensity distribution in phase-space and time. It is applied as a shape function to define the Eddington tensor for the presented VEF model. The shape function computed via the auxiliary RTE problem will capture some degree of transport effects within the TRT problem. The VEF moment equations closed with this approximate Eddington tensor will thus carry with them these captured transport effects. In this study, the temperature data comes from multigroup P 1 , P 1/3 , and flux-limited diffusion radiative transfer models. The proposed VEF model can be interpreted as a transport-corrected diffusion reduced-order model. Numerical results are presented on the Fleck-Cummings test problem which models a supersonic wavefront of radiation. The VEF model is shown to improve accuracy by 1–2 orders of magnitude compared to the considered radiation-diffusion model solutions to the TRT problem.

73 NUCLEAR PHYSICS AND RADIATION PHYSICS↗
NASA NTRS1955

An Investigation of the Effects of Nose and Lip Shapes for an Underslung Scoop Inlet at Mach Numbers from 0 to 1.9

An experimental investigation was conducted to determine the performance characteristics an underslung nose-scoop air-induction system for a supersonic airplane. Five different nose shapes, three lip shapes, and two internal diffusers were investigated. Tests were made at Mach numbers from 0 to 1.9, angles of attack from 0 deg to approximately l5 deg, and mass-flow ratios from 0 to maximum obtainable. It was found that the underslung nose-scoop inlet was able to operate at Mach numbers from 0.6 to 1.9 over a large positive angle-of-attack range without adverse effects on the pressure recovery. Although there was no one inlet configuration that was markedly superior over the entire range of operating variables, the arrangement having a nose designed to give increased supersonic compression at low angles of attack, and a sharp lip (configuration designated N3L3) showed the most favorable performance characteristics over the supersonic Mach number range. Inlets with sizable lip radii gave satisfactory performance up to a Mach number of 1.5; however, as a result of an increase in drag, the performance of such inlets was markedly inferior to the sharp-lip configuration above Mach numbers of 1.5. Throughout the range of test Mach numbers all inlet configurations evidenced stable air-flow characteristics over the mass-flow range for normal engine operation. Analysis of the inlet performance on the basis of a propulsive thrust parameter showed that a fixed inlet area could be used for Mach numbers up to 1.5 with only a small sacrifice in performance.

NOSE CONE↗
NASA NTRS1950

High-Subsonic Performance Characteristics and Boundary-Layer Investigations of a 12 10-Inch-Inlet-Diameter Conical Diffuser

Performance and boundary-layer data were taken in a 12 degree 10-inch inlet-diameter conical diffuser of 2:1 exit- to inlet-area ratio. These data were taken for two inlet-boundary-layer conditions. The first condition was that of a thinner inlet boundary later (boundary-layer displacement thickness, delta* approximately equal to 0.034) produced by an inlet section approximately 1 inlet diameter in length between the entrance bell and the diffuser. The second condition was a thicker inlet boundary layer (delta* approximately equal to 0.120) produced by an additional inlet section length of approximately 6 diameters. Longitudinal static-pressure distributions were measured fro wall static orifices. Transverse total- and static-pressure surveys were made at the inlet and exit stations. Boundary-layer velocity distributions were measured at seven stations between the inlet and exit. These data were obtained for a Reynolds number (based on inlet diameter) range of 1 x 10(exp 6) to 3.9 x 10(exp 6). The corresponding Mach number range was from M = 0.2 to choking. At the maximum-power-available condition supersonic flow was obtained as far as 4.5 inches downstream from the diffuser inlet with a maximum Mach number of M approximately equal to 1.5. The total-pressure loss through the diffuser in percentage of inlet dynamic pressure was approximately 2.5 percent for the thinner inlet boundary later and 5.5 percent for the thicker inlet boundary later over the lower subsonic range. These valued increased with increasing flow rate- the values for the thicker inlet boundary later more than those for the thinner inlet boundary layer. The diffuser effectiveness, expressed as the ratio of the actual static-pressure rise to the ideal static-pressure rise, was about 85 percent for the thinner inlet boundary layer and about 67 percent for the thicker inlet boundary later in the lower subsonic range. These values decrease with increasing flow rate. Separated flow was observed for both inlet-boundary-layer conditions in the region of adverse pressure gradient just downstream of the transition curvature from inlet section to diffuser. The flow for the thinner-inlet-boundary-layer condition did not fully re-establish itself along the diffuser walls. The thicker inlet-boundary-layer flow, while not completely re-establishing the normal flow pattern downstream of the separated region, did re-establish more successfully than the thinner inlet boundary layer.

Little, B. H., Jr.↗
NASA NTRS1992

Parabolized reduced Navier-Stokes computational techniques

A review is presented of methods in which composite or reduced Navier-Stokes (RNS) equations are treated with a pressure-gradient-based flux-vector splitting. The methods are similar to large Re asymptotic formulations, and streamwise diffusion terms are ignored in favor of an explicit deferred corrector based on higher-order diffusion terms. The methods can be used for 2D and 3D supersonic flows in which the effects of real gas are incorporated. Several examples of the procedure are given, and subsonic and supersonic flows are handled well with relaxation procedures that incorporate multigrid acceleration. Effective solutions are described for problems ranging from incompressible flows and supersonic flows to sharp shocks and reverse-flow capturing.

Rubin, Stanley G.↗
DOE OSTI2022

The Eel Pulsar Wind Nebula: A PeVatron-candidate Origin for HAWC J1826-128 and HESS J182-130

HAWC J1826–128 is one of the brightest Galactic TeV γ-ray sources detected by the High Altitude Water Cherenkov (HAWC) observatory, with photon energies extending up to nearly ~100 TeV. This HAWC source spatially coincides with the H.E.S.S. TeV source HESS J1826–130 and the "Eel" pulsar wind nebula (PWN), which is associated with the GeV pulsar PSR J1826–1256. In the X-ray band, Chandra and XMM-Newton revealed that the Eel PWN is composed of both a compact nebula (~15'') and diffuse X-ray emission (~6' × 2') extending away from the pulsar. Our NuSTAR observation detected hard X-ray emission from the compact PWN up to ~20 keV and evidence of the synchrotron burn-off effect. In addition to the spatial coincidence between HESS J1826–130 and the diffuse X-ray PWN, our multiwavelength spectral energy distribution (SED) analysis using X-ray and γ-ray data establishes a leptonic origin of the TeV emission associated with the Eel PWN. Furthermore, our evolutionary PWN SED model suggests (1) a low PWN B-field of ~1 μG, (2) a significantly younger pulsar age (t ~ 5.7 kyr) than the characteristic age (τ = 14.4 kyr), and (3) a maximum electron energy of ${E}_{\max }=2$ PeV. The low B-field, as well as the putative supersonic motion of the pulsar, may account for the asymmetric morphology of the diffuse X-ray emission. Our results suggest that the Eel PWN may be a leptonic PeVatron particle accelerator powered by the ~6 kyr old pulsar PSR J1826–1256 with a spin-down power of 3.6 × 10 36 erg s –1 .

79 ASTRONOMY AND ASTROPHYSICS↗
NASA NTRS1988

Effect of heat release on the spatial stability of a supersonic reacting mixing layer

A numerical study of the stability of compressible mixing layers in which a diffusion flame is embedded is described. The mean velocity profile has been approximated by a hyperbolic tangent profile and the limit of infinite activation energy taken, which reduces the diffusion flame to a flame sheet. The addition of combustion in the form of a flame sheet was found to have important, and complex, effects on the flow stability.

Jackson, T. L.↗
NASA NTRS1964

Review of Problems in Application of Supersonic Combustion

The problem of air-breathing engines capable of flying at very high Mach numbers is described briefly. Possible performance of supersonic combustion ramjets is outlined briefly and the supersonic combustion process is described. Two mechanisms of combustion are outlined: one is supersonic combustion controlled by convection process, and the second is controlled by diffusion. The parameters related to the combustion process are discussed in detail. Data and analyses of reaction rates and mixing phenomena are represented; the flame mechanism is discussed, and experimental results are presented.

Antonio Ferri↗
NASA NTRS1990

Control of flow separation and mixing by aerodynamic excitation

The recent research in the control of shear flows using unsteady aerodynamic excitation conducted at the NASA Lewis Research Center is reviewed. The program is of a fundamental nature, concentrating on the physics of the unsteady aerodynamic processes. This field of research is a fairly new development with great promise in the areas of enhanced mixing and flow separation control. Enhanced mixing research includes influence of core turbulence, forced pairing of coherent structures, and saturation of mixing enhancement. Separation flow control studies included are for a two-dimensional diffuser, conical diffusers, and single airfoils. Ultimate applications include aircraft engine inlet flow control at high angle of attack, wide angle diffusers, highly loaded airfoils as in turbomachinery, and ejector/suppressor nozzles for the supersonic transport. An argument involving the Coanda Effect is made that all of the above mentioned application areas really only involve forms of shear layer mixing enhancement. The program also includes the development of practical excitation devices which might be used in aircraft applications.

Rice, Edward J.↗
NASA NTRS1990

Control of flow separation and mixing by aerodynamic excitation

The recent research progress in the control of shear flows using unsteady aerodynamic excitation conducted at the NASA Lewis Research Center is reviewed. The program is of fundamental nature concentrating on the physics of the unsteady aerodynamic processes. This field of research is a fairly new development with great promise in the areas of enhanced mixing and flow separation control. Enhanced mixing research reported in this paper include influence of core turbulence, forced pairing of coherent structures, and saturation of mixing enhancement. Separation flow control studies included are for a two-dimensional diffuser, conical diffusers, and single airfoils. Ultimate applications of this research include aircraft engine inlet flow control at high angle of attack, wide angle diffusers, highly loaded airfoils as in turbomachinery, and ejector/suppressor nozzles for the supersonic transport. An argument involving the Coanda Effect is made here that all of the above mentioned application areas really only involve forms of shear layer mixing enhancement. The program also includes the development of practical excitation devices which might be used in aircraft applications.

Rice, Edward J.↗
NASA NTRS1994

An Interactive, Design and Educational Tool for Supersonic External-Compression Inlets

A workstation-based interactive design tool called VU-INLET was developed for the inviscid flow in rectangular, supersonic, external-compression inlets. VU-INLET solves for the flow conditions from free stream, through the supersonic compression ramps, across the terminal normal shock region and the subsonic diffuser to the engine face. It calculates the shock locations, the capture streamtube, and the additive drag of the inlet. The inlet geometry can be modified using a graphical user interface and the new flow conditions recalculated interactively. Free stream conditions and engine airflow can also be interactively varied and off-design performance evaluated. Flow results from VU-INLET can be saved to a file for a permanent record, and a series of help screens make the simulator easy to learn and use. This paper will detail the underlying assumptions of the models and the numerical methods used in the simulator.

Benson, Thomas J.↗
NASA NTRS2002

Verification Assessment of Flow Boundary Conditions for CFD Analysis of Supersonic Inlet Flows

Boundary conditions for subsonic inflow, bleed, and subsonic outflow as implemented into the WIND CFD code are assessed with respect to verification for steady and unsteady flows associated with supersonic inlets. Verification procedures include grid convergence studies and comparisons to analytical data. The objective is to examine errors, limitations, capabilities, and behavior of the boundary conditions. Computational studies were performed on configurations derived from a "parameterized" supersonic inlet. These include steady supersonic flows with normal and oblique shocks, steady subsonic flow in a diffuser, and unsteady flow with the propagation and reflection of an acoustic disturbance.

Slater, John W.↗
NASA NTRS1979

Investigation of the feasibility of CARS measurements in scramjet combustion

Results are presented of analytical and experimental investigations to determine the feasibility of using coherent anti-Stokes Raman spectroscopy (CARS) to measure temperature and species concentrations in supersonic combustion experiments. Experimental CARS spectra of O2 and H2 in a laboratory flat H2 - air diffusion flame are presented. Temperatures deduced from both species generally agree with radiation corrected thermocouple measurements. Oxygen concentrations were determined from the shape of the experimental spectra. Measured H2 CARS signal magnitudes are in good agreement with analytical predictions based on H2 concentration profiles determined from spontaneous Raman scattering. Based on these results, CARS diagnostics appear to be feasible if problems associated with the supersonic flow are not too severe.

Shirley, J. A.↗
NASA NTRS1993

Numerical investigations in three-dimensional internal flows

In the present reporting period, the 3D version of the OVERFLOW code was used to solve the flow within the internal portion of the supersonic inlet. The internal portion of this inlet is bounded by an inflow plane containing the leading edge of the sidewalls, the sidewalls, the ramp and cowl surfaces and an outflow plane just downstream of the minimum geometric area of the inlet. Boundary layer bleed was used in the two-dimensional calculations discussed in the previous progress report and that same bleed was applied in the present study. For reference, this bleed corresponds to locations designated as R2 and R3 in the Mach 5 inlet model test. Using the GRIDGEN code, a three dimensional grid was generated that accounted for the viscous effects expected to occur on the sidewall, as well as those known to occur on the ramp and cowl surfaces. The internal flow grid size was 141 streamwise by 101 cross stream by 71 in the lateral direction between sidewalls. Since the flow entering the inlet was not symmetrical, the inlet was solved from sidewall to sidewall (without using a symmetry plane). In addition to the short sidewalls proposed in the Langley geometry database, a set of shorter sidewalls was also investigated in the present study and was shown to have beneficial effects with respect to the flow distortion exiting the supersonic inlet. In addition to these calculations, additional 3D solutions using the OVERFLOW code were obtained for the flow downstream of the throat of the supersonic inlet, including a terminal shock wave system produced by a backpressured subsonic diffuser.

Rose, William C.↗