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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Measurements and Simulations of Surface Dielectric Barrier Discharges Used as Plasma Actuators

This report is a Ph.D. dissertation performed under NRA cooperative agreement and submitted as part of the final report. Asymmetric surface dielectric barrier discharges (DBDs) have shown promise for use as aerodynamic actuators for active flow control. In this project we studied DBD actuators experimentally and numerically. Our DBDs used a symmetric triangular high voltage waveform to generate plasma in atmospheric pressure air. Time-averaged measurements indicated that the induced force of a single barrier actuator design (one electrode insulated from the plasma) can be increased exponentially above the results of previous studies by decreasing both the length and thickness of the electrode exposed to the plasma. This increased force may allow these devices to control flow separation in a wider range of flow environments. Experiments using an intensified digital camera to examine the plasma on time scales of a few nanoseconds showed that, in addition to the previously-observed filamentary and jet-like plasma structures, discharges with very thin exposed electrodes exhibited a weak but constant plasma immediately adjacent to those electrodes. In double-barrier actuators (both electrodes insulated), decreasing the diameter of the narrower electrode lead to increasing forces, and recorded images showed the simultaneous existence of both filamentary and jet-like plasma structures. The development and application of a time-dependent, two-dimensional computational fluid plasma model has aided in understanding the detailed physics of surface DBDs at all-time scales. For simulated single-barrier discharges, the model qualitatively reproduced the filamentary and jet-like micro-discharge structures. The model was somewhat successful in reproducing the observed characteristics of double-barrier actuators. For both actuator geometries, the model indicated that the majority of the forces induced on the neutral gas occur in between micro-discharges as the plasmas decay.

Hoskinson, Alan R.↗
NASA NTRS1985

An incremental block-line-Gauss-Seidel method for the Navier-Stokes equations

A block-line-Gauss-Seidel (LGS) method is developed for solving the incompressible and compressible Navier-Stokes equations in two dimensions. The method requires only one block-tridiagonal solution process per iteration and is consequently faster per step than the linearized block-ADI methods. Results are presented for both incompressible and compressible separated flows: in all cases the proposed block-LGS method is more efficient than the block-ADI methods. Furthermore, for high Reynolds number weakly separated incompressible flow in a channel, which proved to be an impossible task for a block-ADI method, solutions have been obtained very efficiently by the new scheme.

Napolitano, M.↗
NASA NTRSDate not supplied

Three-Dimensional Structure and Modeling of a Normal Bifurcated Shock Train from Experimental Measurements

An asymmetric, bifurcated normal shock train in a Mach 2 constant area, rectangular duct is investigated in the University of Michigan Direct Connect Isolator facility. High-speed schlieren imaging, wall static pressure measurements, surface oil flow visualization, and particle image velocimetry of a shock train are synthesized into a three-dimensional representation of shock train structure. This visualization is then used to inform the underlying flow physics of the distributed fluid dynamical processes along the structure. The detailed, three-dimensional morphology of the flow profile entering the shock train is shown to have a significant impact on the separated flow morphology of the shock train. This results in skewed supersonic core flow with alternating separation bubbles in the surrounding boundary layers. Finally, the efficacy of the prevailing pseudo-shock models in the literature is analyzed with the support of the available flow measurements. None are found to simultaneously model both the pressure and streamwise Mach number profile of the UMDCI pseudo-shock.

Hypersonics↗
NASA NTRSDate not supplied

Three-Dimensional Structure and Modeling of a Normal Bifurcated Shock Train from Experimental Measurements

An asymmetric, normal bifurcated shock train in a Mach 2 constant area, rectangular duct is investigated in the University of Michigan Direct Connect Isolator facility. High-speed schlieren imaging, wall static pressure measurements, surface oil flow visualization, and particle image velocimetry of the shock train are synthesized into a three-dimensional representation. This visualization is then used to inform the underlying flow physics of the distributed fluid dynamical processes along the shock train. The detailed, three-dimensional morphology of the flow profile entering the shock train is shown to have a significant impact on the separated flow morphology of the shock train. This results in a skewed supersonic core flow with alternating separation bubbles in the surrounding boundary layers. Finally, the efficacy of the prevailing pseudo-shock models in the literature is analyzed with the support of the available flow measurements. None are found to model both the pressure and streamwise Mach number profile of the UMDCI pseudo-shock. Note: This presentation is accompanied by the video presentation that can be downloaded for viewing and is formatted as an mp4 with a runtime of 10 min 15 secs.

Hypersonics↗
NASA NTRS1975

Computation of separated transonic turbulent flows

The two-dimensional Reynolds-averaged compressible Navier-Stokes equations are solved using MacCormack's second-order-accurate explicit finite difference method to simulate the separated transonic turbulent flow field over an airfoil. Four different algebraic eddy viscosity models are tested for viability to achieve turbulence closure for the class of flows considered. These models range from an unmodified boundary-layer mixing-length model to a relaxation model incorporating special considerations for the separation bubble region. Results of this study indicate the necessity for special attention to the separated flow region and suggest limits of applicability of algebraic turbulence models to these separated flow fields.

Deiwert, G. S.↗
NASA NTRS2018

Heat Transfer Measurements on the Endwall of a Variable Speed Power Turbine Blade Cascade

Heat transfer measurements were obtained on the endwall of a 2-D section of a variable speed power turbine (VSPT) rotor blade linear cascade. Infrared thermography was used to help determine the transition of flow from laminar to turbulent as well as determine regions of flow separation. Steady state data was obtained for six incidence angles ranging from +15.8 deg to -51 deg, and at five flow conditions for each angle. Nusselt number was used as a method to visualize flow transition and separation on the endwall surface and showed the effects of secondary flows on the surface. Nusselt correlation with Reynolds number from multiple flow conditions was used to plot local values of the correlation exponent and indicated the state of the local boundary layer as the flow transitioned from laminar to turbulent as well as secondary flow features.

variable speed power turbine↗
NASA NTRSDate not supplied

In-Tunnel Simulations of the NASA Juncture Flow Model

The effect of wind tunnel walls on the development of flow separation in the wing-fuselage juncture area is assessed using FUN3D’s overset and wind tunnel controller capabilities. The simulation data obtained from the in-tunnel runs are compared with the free-air simulations. The simulation results are also evaluated using measurements from the recently completed Phase 2 of the Juncture Flow Experiment. In general, the simulation data compared well with the measurements, however there were no significant differences in the prediction of the separation location, size, and shape, between the free-air and the in-tunnel runs. The results of this study suggest that the effects of tunnel walls on the flow separation in the juncture region are relatively small.

Juncture Flow↗
NASA NTRSDate not supplied

In-Tunnel Simulations of the NASA Juncture Flow Model

The effect of wind tunnel walls on the development of flow separation in the wing-fuselage juncture area is assessed using FUN3D’s overset and wind tunnel controller capabilities. The simulation data obtained from the in-tunnel runs are compared with the free-air simulations. The simulation results are also evaluated using measurements from the recently completed Phase 2 of the Juncture Flow Experiment. In general, the simulation data compared well with the measurements, however there were no significant differences in the prediction of the separation location, size, and shape, between the free-air and the in-tunnel runs. The results of this study suggest that the effects of tunnel walls on the flow separation in the juncture region are relatively small.

Juncture Flow↗
NASA NTRS1985

Symposium on Numerical and Physical Aspects of Aerodynamic Flows, 3rd, California State University, Long Beach, CA, January 21-24, 1985, Proceedings

Among the topics discussed are hydrodynamic stability and turbulent transition, separated flow simulations employing finite difference methods, a quasi-simultaneous finite difference approach for strongly interacting flow, the Newton solution of coupled Euler and boundary layer equations, multielement airfoil aerodynamics, vorticity dynamics, the significance of the thin layer Navier-Stokes approximation, a comparison of interactive and Navier-Stokes calculations of separating boundary layer flows, and a Navier-Stokes analysis for the mechanism of shock wave/boundary layer interactions. Also discussed are transonic interactions of unsteady vortical flows, boundary layers on oscillating airfoils, a laser-interferometer skin friction meter, flow over surface-mounted semibluff bodies, computations for spinning and spin-stabilized supersonic projectiles, transition calculations in three-dimensional flows, the effect of projectile boattailing in transonic flow, flow characteristics of a body of revolution at incidence, the influence of the boundary layer on the leading edge flow of a swept wing, and an interactive scheme for three-dimensional transonic flows.

Source record↗
NASA NTRS1982

An experimental documentation of a separated trailing-edge flow at a transonic Mach number

A detailed experiment on the separated flow field at a sharp trailing edge is described and documented. The separated flow is a result of sustained adverse pressure gradients. The experiment was conducted using an elongated airfoil-like model at a transonic Mach number and at a high Reynolds number of practical interest. Measurements made include surface pressures and detailed mean and turbulence flow quantities in the region just upstream of separation to downstream into the near-wake, following wake closure. The data obtained are presented mostly in tabular form. These data are of sufficient quality and detail to be useful as a test case for evaluating turbulence models and calculation methods.

Viswanath, P. R.↗
NASA NTRS1993

Numerical simulation of a shock wave/turbulent boundary layer interaction in a duct

A numerical investigation of the interaction of an incident oblique shock wave with a turbulent duct flow is presented. The investigation consists of solving the three-dimensional, unsteady, compressible, mass averaged Navier-Stokes equations, using an implicit finite volume, lower-upper time marching code and incorporates the three-dimensional Baldwin-Lomax turbulence model. Computed results are obtained Mach number 2.9 for a turning angle of 13 degrees and Reynolds number based on duct width of 1.36 x 10 exp 7. Under various inlet conditions, the results clearly depict the flow characteristics, including the shock geometry, the separated flow region, the wall pressure distribution, and the skin friction distribution. The findings provide a physical understanding of the three-dimensional vortex structure of the flow in a duct in which a shock wave interacts with a turbulent boundary layer. The results show that the ratio of the boundary layer thickness to the duct width is the critical parameter in determining the separation structure.

Yang, Wei-Li↗
NASA NTRS2002

Active Control of Separation from the Slat Shoulder of a Supercritical Airfoil

Active flow control in the form of zero-mass-flux excitation was applied at the slat shoulder of a simplified high-lift airfoil to delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge slat and a 25% chord simply hinged trailing edge flap. The cruise configuration data was successfully reproduced, repeating previous experiments. The effects of flap and slat deflection angles on the performance of the airfoil integral parameters were quantified. Detailed flow features were measured as well, in an attempt to identify optimal actuator placement. The measurements included: steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization and Particle Image Velocimetry (PIV). High frequency periodic excitation was applied to delay the occurrence of slat stall and improve the maximum lift by 10 to 15%. Low frequency amplitude modulation was used to reduce the oscillatory momentum coefficient by roughly 50% with similar aerodynamic performance.

Pack, LaTunia G.↗
NASA NTRSDate not supplied

HLPW-4/GMGW-3: Wall-Modeled LES and Lattice-Boltzmann Technology Focus Group Workshop Summary

A summary of the nine submissions to the Wall-Modeled LES and Lattice-Boltzmann(WMLESLB) Technical Focus Group (TFG) at the 4th High lift Prediction Workshop is provided. The focus of this TFG was to assess the current capabilities of WMLES and Lattice-Boltzmann methods on a complex high-lift configuration across a wide range of angles of attack. Preliminary analysis of the submitted data suggests that>250𝑀spatial degrees of freedom are needed to accurately predict pitching moments at high angles of attack due to large pressure gradients present on the outboard slat and main element for𝛼 >17◦(corrected for free-air).While some scatter is reported in pitching moments at the low-angles of attack (𝛼 <11◦) - likely caused by differences in flap separation possibly due to low Reynolds number effects - excellent agreement is observed between the submissions near the𝐶𝐿,max state. Objective superiority over steady state RANS can be seen in terms of lack of excess outboard separation; a majority of the good quality WMLES and LB submissions predict wedge-shaped separation patterns consistent with the experimental oil flow. Differences in the onset of stall mechanism in the free-air configuration for𝛼 >20◦is reported with two distinct topologies observed. Topology A is characterized by the onset of corner-flow separation which progressively grows to produce a pitch break in free-air with an angle of attack increase from𝛼=20.55◦to𝛼=21.47◦+𝜖where𝜖is a small perturbation (varying between submissions and likely to be within±0.3◦). Topology Bis characterized by boundary layer weakness emanating from the inboard side of the wing-pylon juncture substantially larger than any weakness in the wing-body juncture (due to corner-flow);submissions in this category do not show any tendency for occurrence of a pitch break for the free-air configuration within the vicinity of𝛼=21.47◦. The in-tunnel simulations submitted by3 participants using different discretizations, grids and closure models show excellent agreement with the experiment in terms of a) integrated loads, b) surface flow-topology, and c) mechanism for the onset of inboard stall. Further evidence is provided to demonstrate both qualitative and quantitative superiority of all 3 WMLES submissions over a single steady state RANS submission to the workshop.

TTT↗
NASA NTRS2019

Turbulence Measurements on a Wing-Fuselage Junction Model for CFD Validation

Current turbulence models, such as those employed in Reynolds-averaged Navier-Stokes CFD, are unable to reliably predict the onset and extent of the three-dimensional separated flow that typically occurs in wing fuselage junctions. To critically assess, as well as to improve upon, existing turbulence models, experimental validation-quality flow field data in the junction region is needed. To that end, NASA has developed a CFD validation experiment for a generic full-span wing-fuselage junction model at subsonic conditions. In this paper, we present and discuss sample results from this effort, with a focus on turbulence measurements in the separated corner flow of the wing-fuselage junction that were obtained with an internally mounted laser Doppler velocimetry (LDV) system. In addition, sample comparisons between the flow field measurements and representative CFD of our test case are presented and here, we highlight some of the key discrepancies.

Kegerise, Michael↗
DOE OSTI2013

Petascale Adaptive CFD (ALCF ESP Technical Report)

The aerodynamic simulations of this project involve modeling of active flow control based on synthetic jet actuation that has been shown experimentally to produce large-scale flow changes (e.g., re-attachment of separated flow or virtual aerodynamic shaping of lifting surfaces) from micro-scale input. This micro-scale input consists for instance in 0.1 W piezoelectric disk which resonate in a cavity alternately and pushes/pulls out/in the fluid through a small slit to create small-scale vortical structures that interact with, and thereby dramatically alter, the cross flow. This process has yet to be understood fundamentally. Synthetic jet actuators offer the prospect of not only augmenting lift but also other forces and moments in a dynamic and controlled fashion for a range of operating conditions. They have been demonstrated to restore and maintain flow attachment and reduce vibrations in wind turbine blades during dynamic pitch, thereby reducing unsteady loads on gearboxes that are currently the prime failure point. In virtual-shape flight control surfaces for aerial vehicles (including commercial airplanes), conventional control surfaces (e.g., flaps, rudder, etc.) can be augmented or even replaced with active flow control, thus improving their lift-to-drag ratio and/or control power. The goal of the numerical simulations proposed in this project is to provide a complementary and detailed view of the flow interactions at a much higher Reynolds number, approaching engineering application scales for the first time.

Jansen, Kenneth E.↗