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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 1,009 records · Page 56
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Advanced Subsonic Technology (AST) Area of Interest (AOI) 6: Develop and Validate Aeroelastic Codes for Turbomachinery

AlliedSignal Engines, in cooperation with NASA GRC (National Aeronautics and Space Administration Glenn Research Center), completed an evaluation of recently-developed aeroelastic computer codes using test cases from the AlliedSignal Engines fan blisk and turbine databases. Test data included strain gage, performance, and steady-state pressure information obtained for conditions where synchronous or flutter vibratory conditions were found to occur. Aeroelastic codes evaluated included quasi 3-D UNSFLO (MIT Developed/AE Modified, Quasi 3-D Aeroelastic Computer Code), 2-D FREPS (NASA-Developed Forced Response Prediction System Aeroelastic Computer Code), and 3-D TURBO-AE (NASA/Mississippi State University Developed 3-D Aeroelastic Computer Code). Unsteady pressure predictions for the turbine test case were used to evaluate the forced response prediction capabilities of each of the three aeroelastic codes. Additionally, one of the fan flutter cases was evaluated using TURBO-AE. The UNSFLO and FREPS evaluation predictions showed good agreement with the experimental test data trends, but quantitative improvements are needed. UNSFLO over-predicted turbine blade response reductions, while FREPS under-predicted them. The inviscid TURBO-AE turbine analysis predicted no discernible blade response reduction, indicating the necessity of including viscous effects for this test case. For the TURBO-AE fan blisk test case, significant effort was expended getting the viscous version of the code to give converged steady flow solutions for the transonic flow conditions. Once converged, the steady solutions provided an excellent match with test data and the calibrated DAWES (AlliedSignal 3-D Viscous Steady Flow CFD Solver). However, efforts expended establishing quality steady-state solutions prevented exercising the unsteady portion of the TURBO-AE code during the present program. AlliedSignal recommends that unsteady pressure measurement data be obtained for both test cases examined for use in aeroelastic code validation.

Gardner, Kevin D.↗
DOE OSTI2021

An exact inversion method for extracting orientation ordering by small-angle scattering

Here, we outline a nonparametric inversion strategy for determining the orientation distribution function (ODF) of sheared interacting rods using small-angle scattering techniques. With the presence of direct inter-rod interaction and fluid mechanical forces, the scattering spectra are no longer characterized by the azimuthal symmetry in the coordinates defined by the principal directions of simple shear conditions, which severely compounds the reconstruction of ODFs based on currently available methods developed for dilute systems. Using a real spherical harmonic expansion scheme, the real-space ODFs are uniquely determined from the anisotropic scattering spectra and their numerical accuracy is verified computationally. Our method can be generalized to extract ODFs of uniaxially anisotropic objects under different flow conditions in a properly transformed reference frame with suitable basis vectors.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
DOE OSTI2020

Mesoscale to Microscale Simulations over Complex Terrain with the Immersed Boundary Method in the Weather Research and Forecasting Model

Improvements to the Weather Research and Forecasting (WRF) Model are made to enable multiscale simulations over highly complex terrain with dynamically downscaled boundary conditions from the mesoscale to the microscale. Over steep terrain, the WRF Model develops numerical errors that are due to grid deformation of the terrain-following coordinates. An alternative coordinate system, the immersed boundary method (IBM), has been implemented into WRF, allowing for simulations over highly complex terrain; however, the new coordinate system precluded nesting within mesoscale simulations using WRF’s native terrain-following coordinates. Here, the immersed boundary method and WRF’s grid-nesting framework are modified to seamlessly work together. This improved framework for the first time allows for large-eddy simulation over complex (urban) terrain with IBM to be nested within a typical mesoscale WRF simulation. Simulations of the Joint Urban 2003 field campaign in Oklahoma City, Oklahoma, are performed using a multiscale five-domain nested configuration, spanning horizontal grid resolutions from 6 km to 2 m. These are compared with microscale-only simulations with idealized lateral boundary conditions and with observations of wind speed/direction and SF 6 concentrations from a controlled release from intensive observation period 3. The multiscale simulation, which is configured independent of local observations, shows similar model skill predicting wind speed/direction and improved skill predicting SF 6 concentrations when compared with the idealized simulations, which require use of observations to set mean flow conditions. Use of this improved multiscale framework shows promise for enabling large-eddy simulation over highly complex terrain with dynamically downscaled boundary conditions from mesoscale models.

42 ENGINEERING↗
NASA NTRS1993

Viscous hypersonic shock-shock interaction on a blunt body at high altitude

The shock interaction produced when an incident oblique shock impinges on a blunt body, such as an engine inlet cowl lip of a hypersonic vehicle, is investigated for high altitude flight conditions. A perfect gas, Navier-Stokes numerical simulation of this problem at various altitudes representing continuum through transitional conditions is performed using the modified flux vector splitting method of Steger and Warming (1979). Two series of solutions are produced. First, a number-of-shock-positions are studied at a particular altitude and Mach number. Second, given a fixed shock position and Mach number, the interaction is investigated at several altitudes ranging from continuum to transitional flow conditions. It is shown that the interaction becomes fundamentally different as the density is lowered, and its effect on the overheating problem is progressively diminished. The maximum stagnation point heating at the highest altitude is reached only when the incident shock misses the cowl lip completely, and any interaction with the cowl bow shock that does occur takes place downstream and thus has little effect on the conditions at the stagnation point.

Comeaux, Keith A.↗
NASA NTRS1995

Effects of Chemistry on Blunt-Body Wake Structure

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

Dogra, Virendra K.↗
DOE OSTI2026

A filter-dependent granular temperature model from large-scale CFD-DEM data

The computational study of strongly-coupled, gas–solid flows at scales relevant to most environmental and engineering applications requires the use of ‘coarse-grained’ methodologies such as the two-fluid model, particle-in-cell approach or the multiphase Reynolds Averaged Navier–Stokes equations. While these strategies enable computations at desirable length- and time-scales, they rely heavily on models to capture important flow physics that occur at scales smaller than the mesh. To date, the models that do exist are based on a limited set of flow conditions, such as very dilute particle phase. To this end, we leverage a large-scale repository of CFD-DEM data to develop filter-size dependent models for the mean variance in particle volume fraction, a quantity commonly used to assess the degree of clustering, and the granular temperature, a key quantity for accurately predicting gas–solid flows. In conclusion, because of its filter-size dependence, the granular temperature model can be directly translated to coarse-grained approaches and tied directly to grid size.

AMReX↗
NASA NTRS1977

Numerical solution of the viscous hypersonic flow past blunted cones at angle of attack

Hypersonic viscous flow over spherically blunted cones of large half angle is computed at small angles of attack in the plane of symmetry of the flow field. Time-dependent viscous shock-layer equations in body-oriented coordinates are used to describe the flow field. The shock wave is treated as a discontinuity, across which the Rankine-Hugoniot relations are used to compute the flow conditions behind the shock. A time-marching second-order finite-difference method is used to solve the equations for a perfect gas. The local CFL (Courant-Friedrich-Lewy) time increment is used to advance the solution in time at each grid point. A fourth-order damping is used to damp the oscillations in the flow quantities. The numerical results of the present analysis for quantities such as shock standoff distance, surface-pressure distribution, and heating rates compare well with the existing theoretical and experimental results.

Kumar, A.↗
NASA NTRS1987

Transonic characteristics of a humped airfoil

A humped airfoil concept to improve the perfomance of supercritical airfoils at off-design conditions in transonic maneuver is introduced. Theoretical aspects of the airfoil and recent results of experimental tests conducted in the high Reynolds number transonic wind tunnels at both NASA Ames Research Center and The Ohio State University are presented. Experimental evidence has shown that the humped airfoil generally has more favorable transonic characteristics in off-design high supercritical flow conditions than a regular supercritical airfoil.

Tai, Tsze C.↗
NASA NTRS1997

Combustion Noise at Elevated Pressures in a Liquid-Fueled Premixed Combustor

Noise generated in gas turbine combustors can exist in several forms-broadband noise, sharp resonant peaks, and regular or intermittent nonlinear pulsing. In the present study, dynamic pressure measurements were made in several JP-5-fueled combustor configurations, at various mean pressures and temperatures. The fluctuating pressure was measured at mean pressures from 6 to 14 atm and inlet temperatures from 550 K to 850 K. The goal of the present work was to study the effect of changes in mean flow conditions on combustor noise: both broadband noise and sharp tones were considered. In general, the shape of the broadband noise spectrum was consistent from one configuration to another. The shape of the spectrum was influenced by the acoustic filtering of the combustion zone. This filtering ensured the basic consistency of the spectra. In general, the trends in broadband noise observed at low mean pressures were also seen at high mean pressures; that is, the total sound level decreased with both increasing equivalence ratio and increasing inlet temperature. The combustor configurations without a central pilot experienced higher broadband noise levels and were more susceptible to narrow peak resonances than configurations with a central pilot. The sharp peaks were more sensitive to the mean flow than was the broadband noise, and the effects were not always the same. In some situations, increasing the equivalence ratio made the sharp peaks grow, while at other conditions, increasing the equivalence ratio made the sharp peaks shrink. Thus, it was difficult to predict when resonances would occur; however, they were reproducible. Acoustic coupling between the upstream and downstream regions of the combustor may play a role in the sharp-peaked oscillations. Noise was also observed near lean blow out. As with other types of noise, lean blow out noise was affected by the combustion chamber acoustics, which apparently maintains the fluctuations at a uniform frequency. However, the actual conditions when this type of noise was experienced appeared to simply follow the lean blow out limit as it varied with mean temperature and pressure.

Darling, Douglas↗
NASA NTRS1991

A numerical study of shock wave/boundary layer interaction in nonequilibrium chemically reacting air - The effects of catalytic walls

This paper presents a numerical study that investigate the effects of nonequilibrium chemistry, and in particular, wall catalycity on the separated flow region generated by an oblique shock wave impinging upon a flat plate boundary layer. To obtain a solution to this problem, the full two dimensional Navier-Stokes equations were solved using MacCormack's predictor-corrector time dependent technique on a rectangular grid. Nonequilibrium chemistry was included by utilizing the 5 species, 17 reaction modified Dunn-Kang chemical kinetics model. Separate results were obtained for: calorically perfect, chemically reacting - noncatalytic wall, and chemically reacting - fully catalytic wall cases, for a given set of flow conditions. A direct comparison of all three cases revealed a slight decrease in the peak heat transfer for the noncatalytic wall case, as compared to the calorically perfect case. On the other hand, the fully catalytic wall case had a tremendous increase in the peak surface heat transfer. It is concluded that, for the particular conditons treated here (nearly frozen flow in the free stream), the effects of the nonequilibrium chemically reacting flow on the shock-wave/boundary-layer interaction depend critically on the catalycity of the wall, having virtually no effect for the case of a noncatalytic wall, and exerting a tremendous effect for a fully catalytic wall.

Grumet, Adam A.↗
NASA NTRS2002

A microarray immunoassay for simultaneous detection of proteins and bacteria

We report the development and characterization of an antibody microarray biosensor for the rapid detection of both protein and bacterial analytes under flow conditions. Using a noncontact microarray printer, biotinylated capture antibodies were immobilized at discrete locations on the surface of an avidin-coated glass microscope slide. Preservation of capture antibody function during the deposition process was accomplished with the use of a low-salt buffer containing sucrose and bovine serum albumin. The slide was fitted with a six-channel flow module that conducted analyte-containing solutions over the array of capture antibody microspots. Detection of bound analyte was subsequently achieved using fluorescent tracer antibodies. The pattern of fluorescent complexes was interrogated using a scanning confocal microscope equipped with a 635-nm laser. This microarray system was employed to detect protein and bacterial analytes both individually and in samples containing mixtures of analytes. Assays were completed in 15 min, and detection of cholera toxin, staphylococcal enterotoxin B, ricin, and Bacillus globigii was demonstrated at levels as low as 8 ng/mL, 4 ng/mL, 10 ng/mL, and 6.2 x 10(4) cfu/mL, respectively. The assays presented here are very fast, as compared to previously published methods for measuring antibody-antigen interactions using microarrays (minutes versus hours).

Non-NASA Center↗
NASA NTRS2022

Stability Analysis of Streaks Induced by Optimized Vortex Generators

Numerical computations are performed to investigate the potential for transition control in an axisymmetric boundary layer via fully realizable, streamwise stationary streaks induced by an azimuthally periodic array of surface mounted vortex generators (VGs). Previous work has shown that suitable streaks of this type can significantly reduce the growth of Mack’s second mode instabilities, but large streak amplitudes can make the flow susceptible to previously absent streak instabilities that can become the leading cause of transition. Here, we use the adjoint capabilities of the SU2 flow solver to optimize the VG shape to maximize the reduction in the growth of second-mode disturbances while also preventing the streak amplitudes from reaching large enough values to precipitate an earlier onset of transition via streak instabilities. The geometry and the freestream flow conditions are selected to match a relevant trajectory lo-cation from the HIFiRE-1 flight experiment. Results show that the optimized VGs can increase the mean streak amplitude by 117% with respect to a manually developed baseline design. The stability of this optimized basic state is analyzed via the plane-marching parabolized stability equations, predicting a fully laminar flow over the entire cone, or equivalently, yielding transition delay of 130% versus the 17% for the baseline VGs.

Boundary layer transition↗
NASA NTRS2022

Stability Analysis of Streaks Induced By Optimized Vortex Generators

Numerical computations are performed to investigate the potential for transition control in an axisymmetric boundary layer via fully realizable, streamwise stationary streaks induced by an azimuthally periodic array of surface mounted vortex generators (VGs). Previous work has shown that suitable streaks of this type can significantly reduce the growth of Mack’s second mode instabilities, but large streak amplitudes can make the flow susceptible to previously absent streak instabilities that can become the leading cause of transition. Here, we use the adjoint capabilities of the SU2 flow solver to optimize the VG shape to maximize the reduction in the growth of second-mode disturbances while also preventing the streak amplitudes from reaching large enough values to precipitate an earlier onset of transition via streak instabilities. The geometry and the freestream flow conditions are selected to match a relevant trajectory lo-cation from the HIFiRE-1 flight experiment. Results show that the optimized VGs can increase the mean streak amplitude by 117% with respect to a manually developed baseline design. The stability of this optimized basic state is analyzed via the plane-marching parabolized stability equations, predicting a fully laminar flow over the entire cone, or equivalently, yielding transition delay of 130% versus the 17% for the baseline VGs.

Boundary layer transition↗
NASA NTRS2019

Measurement and Analysis of Terminal Shock Oscillation and Buffet Forcing Functions on a Launch Vehicle Payload Fairing

The buffet loads on a launch vehicle payload shroud can be impacted by the unsteadiness associated with a terminal shock at high subsonic speeds. At these conditions, flow accelerates to supersonic speeds on the nose of the payload fairing and is terminated by a normal shock on the cylindrical section downstream of the nose cone/cylinder shoulder. The location of the terminal shock and associated separated boundary layer is affected by the freestream Mach number, Reynolds number, and the pitch/yaw of the launch vehicle. Furthermore, even when the freestream conditions and vehicle attitude are constant, this terminal shock oscillates on the surface of the vehicle. The time-varying surface pressure associated with the terminal shock results in unsteady aerodynamic loads that may interact with vehicle structural dynamic modes and the guidance and control of the vehicle. Buffet testing of a 3-percent scale rigid buffet model of a launch vehicle cargo configuration with a tangent-ogive payload shroud was conducted in 2012 and in 2016. Initial buffet forcing functions (BFFs) utilized a coarse pressure sensor distribution on the vehicle surface in which a single longitudinal station with eight sensors observed the terminal shock environment at Mach 0.90. An examination of these circumferential pressures reveal large impulse-like pressure fluctuations and an asymmetry in pressure when the vehicle is at a nonzeroangle of attack that result in high BFFs. Revisions to the shock integration region were made based on computational fluid dynamics and shadowgraph video of shock motion to better represent the BFFs and reduce the high loads resulting from this environment. To more clearly understand this terminal shock environment, a second wind tunnel test was conducted with a dense distribution of 256 sensors at the terminal shock location. These sensor arrays presents a unique opportunity to observe the unsteady terminal shock environment and to characterize the impact of various integration schemes on the BFFs. This paper presents a summary of the development of BFFs for this terminal shock and a detailed analyses of shock region pressure coefficients, coherence, BFFs, shock location time histories, and power spectral density to help guide development of BFFs for other launch vehicle test and analysis programs.

Piatak, David J.↗
NASA NTRS2020

Simulation and Modeling of Cold-Wall Hypersonic Turbulent Boundary Layers on Flat Plate

Direct numerical simulations (DNS) of flat-plate, zero-pressure-gradient turbulent boundary layers are presented for nominal freestream Mach numbers of 11 and 14 and a highly cooled wall (wall-to-recovery temperature of approximately 0.2). The flow conditions of the DNS are representative of the experimental data for a Mach 11.1 turbulent boundary layer on a flat plate that was tested at Calspan–University of Buffalo Research Center (CUBRC) and the operational conditions of the AEDC Hypervelocity Tunnel No. 9 at Mach 14. The wall shear stress and turbulent heat flux predicted by DNS show good comparisons with those measured at CUBRC and those modeled with a Baldwin-Lomax model. The generated DNS database extends the range of Reynolds numbers from those typical of previous DNS studies for hypersonic boundary-layer flows and confirms the validity of Morkovin’s hypothesis within the relatively broader range of Reynolds numbers considered in this study. A priori assessment of turbulent heat flux models with DNS data shows that the algebraic energy flux (AEF) formulation gives better predictions of the streamwise turbulent heat flux than the commonly used constant turbulent Prandtl number model while both models give similarly good predictions of the wall-normal turbulent heat flux. A posteriori assessment shows that the AEF model improves the prediction of the temperature profiles while keeping a good prediction of the velocity field.

Junji Huang↗
NASA NTRS2012

Shock Generation and Control Using DBD Plasma Actuators

This report is the final report of a NASA Phase I SBIR contract, with some revisions to remove company proprietary data. The Shock Boundary Layer Interaction (SBLI) phenomena in a supersonic inlet involve mutual interaction of oblique shocks with boundary layers, forcing the boundary layer to separate from the inlet wall. To improve the inlet efficiency, it is desired to prevent or delay shock-induced boundary layer separation. In this effort, Innovative Technology Applications Company (ITAC), LLC and the University of Notre Dame (UND) jointly investigated the use of dielectric-barrier-discharge (DBD) plasma actuators for control of SBLI in a supersonic inlet. The research investigated the potential for DBD plasma actuators to suppress flow separation caused by a shock in a turbulent boundary layer. The research involved both numerical and experimental investigations of plasma flow control for a few different SBLI configurations: (a) a 12 wedge flow test case at Mach 1.5 (numerical and experimental), (b) an impinging shock test case at Mach 1.5 using an airfoil as a shock generator (numerical and experimental), and (c) a Mach 2.0 nozzle flow case in a simulated 15 X 15 cm wind tunnel with a shock generator (numerical). Numerical studies were performed for all three test cases to examine the feasibility of plasma flow control concepts. These results were used to guide the wind tunnel experiments conducted on the Mach 1.5 12 degree wedge flow (case a) and the Mach 1.5 impinging shock test case (case b) which were at similar flow conditions as the corresponding numerical studies to obtain experimental evidence of plasma control effects for SBLI control. The experiments also generated data that were used in validating the numerical studies for the baseline cases (without plasma actuators). The experiments were conducted in a Mach 1.5 test section in the University of Notre Dame Hessert Laboratory. The simulation results from cases a and b indicated that multiple spanwise actuators in series and at a voltage of 75 kVp-p could fully suppress the flow separation downstream of the shock. The simulation results from case c showed that the streamwise plasma actuators are highly effective in creating pairs of counter-rotating vortices, much like the mechanical vortex generators, and could also potentially have beneficial effects for SBLI control. However, to achieve these effects, the positioning and the quantity of the DBD actuators used must be optimized. The wind tunnel experiments mapped the baseline flow with good agreement to the numerical simulations. The experimental results were conducted with spanwise actuators for cases a and b, but were limited by the inability to generate a sufficiently high voltage due to arcing in the wind-tunnel test-section. The static pressure in the tunnel was lower than the static pressure in an inlet at flight conditions, promoting arching and degrading the actuator performance.

Patel, Mehul P.↗
NASA NTRS1975

Compliant wall-turbulent skin-friction reduction research

Previous compliant-wall experiments successful in reducing skin-friction drag probably have had a (unplanned) membrane resonance at a favorable frequency, amplitude, wave shape, length, and speed. The most probable drag reduction mechanism involves a direct coupling between the fluid and the moving wall when the wall natural resonance frequencies are near the fundamental turbulent burst frequency. Local skin-friction reductions of 61% were measured with mylar/PVC plastisol compliant surfaces. These reductions were observed only at certain flow conditions, indicating that changing tunnel total temperature may have altered the substrate dynamic modulus, damping, and coupled mylar tension. Apparently, the coupled membrane/substrate must be excited in compatible narrow-band natural frequency modes. An accelerated effort is required to develop practical durable compliant surfaces optimized for maximum drag reduction. Application of compliant walls to other transportation modes appears feasible with liquid flows offering the greatest skin-friction drag reduction potential.

Fischer, M. C.↗
NASA NTRS1975

FORTRAN program for predicting off-design performance of radial-inflow turbines

The FORTRAN IV program uses a one-dimensional solution of flow conditions through the turbine along the mean streamline. The program inputs needed are the design-point requirements and turbine geometry. The output includes performance and velocity-diagram parameters over a range of speed and pressure ratio. Computed performance is compared with the experimental data from two radial-inflow turbines and with the performance calculated by a previous computer program. The flow equations, program listing, and input and output for a sample problem are given.

Wasserbauer, C. A.↗