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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 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

The Effects of Gravity on Wrinkled Laminar Flames

The effects of gravity are significant to the dynamics of idealized unconfined open premixed flames. Moderate to low turbulence Reynolds number flames, i.e., wrinkled laminar flames, of various unconfined geometries have been used extensively for investigating fundamental processes of turbulent flame propagation and to validate theoretical models. Without the wall constraints, the flames are free to expand and interact with surrounding ambient air. The flow field in which the flame exists is determined by a coupling of burner geometry, flame orientation and the gravity field. These complex interactions raise serious questions regarding the validity of comparing the experimental data of open flames with current theoretical and numerical models that do not include the effects of gravity nor effects of the larger aerodynamic flowfield. Therefore, studies of wrinkled laminar flame in microgravity are needed for a better understanding of the role of gravity on flame characteristics such as the orientation, mean aerodynamics stretch, flame wrinkle size and burning rate. Our approach to characterize and quantify turbulent flame structures under microgravity is to exploit qualitative and quantitative flow visualization techniques coupled with video recording and computer controlled image analysis technologies. The experiments will be carried out in the 2.2 second drop tower at the NASA Lewis Research Center. The longest time scales of typical wrinkled laminar flames in the geometries considered here are in the order of 10 msec. Hence, the duration of the drop is sufficient to obtain the amount of statistical data necessary for characterize turbulent flame structures.

Kostiuk, Larry W.↗
NASA NTRS1986

Boundary-layer computational model for predicting the flow and heat transfer in sudden expansions

Fully developed turbulent and laminar flows through symmetric planar and axisymmetric expansions with heat transfer were modeled using a finite-difference discretization of the boundary-layer equations. By using the boundary-layer equations to model separated flow in place of the Navier-Stokes equations, computational effort was reduced permitting turbulence modelling studies to be economically carried out. For laminar flow, the reattachment length was well predicted for Reynolds numbers as low as 20 and the details of the trapped eddy were well predicted for Reynolds numbers above 200. For turbulent flows, the Boussinesq assumption was used to express the Reynolds stresses in terms of a turbulent viscosity. Near-wall algebraic turbulence models based on Prandtl's-mixing-length model and the maximum Reynolds shear stress were compared.

Lewis, J. P.↗
NASA NTRS1973

Preliminary lift and drag characteristics of the F-8 supercritical wing airplane

A flight research program is being conducted using a supercritical wing attached to a modified F-8 fuselage to demonstrate the potential benefits of the supercritical wing concept in a realistic cruise flight environment. One of the primary goals of the flight program is to determine the drag-divergence Mach number, that is, the Mach number at which the total vehicle drag increases significantly. Another objective of the program is to show how well wind-tunnel data can be used to predict the performance of a supercritical wing configuration in the transonic Mach number region from 0. 9 to 1. 0. Comparisons of this nature are important in view of the uncertainties in the present methods for extrapolating wind-tunnel-model data to flight Reynolds numbers, determining the effect of wall interference on the model results, and simulating the flow separation characteristics on the model wing to correspond to the large-scale wing. However, in comparing these wind-tunnel and flight results, it must be taken into account that this supercritical wing test configuration is of only intermediate scale und that a full-scale transport would have Reynolds numbers three to five times larger.

Jon S. Pyle↗
NASA NTRS1990

Calculations of gaseous H2/O2 thruster

Calculations were made for the gaseous H2/O2 thruster for the Space Station by using the RPLUS code which employs an implicit finite volume, LUSSOR scheme to solve the Navier-Stokes equations and the species equations. The combustion processes of hydrogen and oxygen are modeled by a 9 species and 18 step reaction mechanism and the turbulence is simulated by the Baldwin-Lomax turbulence model for the thruster wall boundary layer and the modified Prandtl's mixing length model for the reacting shear layer. Results are presented for different mixture ratios and fuel film cooling percents and compared with the experimental data. The calculated performance predictions for the thruster agree well with the experimental data and the results demonstrate that the RPLUS code can be used for design and analysis of thrusters and rockets.

Kim, S. C.↗
NASA NTRS1994

Simulation of a turbulent flame in a channel

The interaction between turbulent premixed flames and channel walls is studied. Combustion is represented by a simple irreversible reaction with a large activation temperature. Feedback to the flowfield is suppressed by invoking a constant density assumption. The effect of wall distance on local and global flame structure is investigated. Quenching distances and maximum wall heat fluxes computed in laminar cases are compared to DNS results. It is found that quenching distances decrease and maximum heat fluxes increase relative to laminar flame values. It is shown that these effects are due to large coherent structures which push flame elements towards to wall. The effect of wall strain is studied in flame-wall interaction in a stagnation line flow; this is used to explain the DNS results. It is also shown that 'remarkable' flame events are produced by interaction with a horseshoe vortex: burnt gases are pushed towards the wall at high speed and induce quenching and high wall heat fluxes while fresh gases are expelled from the wall region and form finger-like structures. Effects of the wall on flame surface density are investigated, and a simple model for flame-wall interaction is proposed; its predictions compare well with the DNS results.

Bruneaux, G.↗
NASA NTRS1998

Thermal Modelling of Various Thermal Barrier Coatings in a High Flux Rocket Engine

A thermal model was developed to predict the thermal response of coated and uncoated tubes tested in a H2/O2 rocket engine. Temperatures were predicted for traditional APS ZrO2-Y2O3 thermal barrier coatings, as well as APS and LPPS ZrO2-Y2O3/NiCrAlY cermet coatings. Good agreement was observed between predicted and measured metal temperatures at locations near the tube surface or at the inner tube wall. The thermal model was also used to quantitatively examine the effect of various coating system parameters on the temperatures in the substrate and coating. Accordingly, the effect of the presence a metallic bond coat and the effect of radiation from the surface of the ceramic layer were examined. In addition, the effect of a variation in the values of the thermal conductivity of the ceramic layer was also investigated. It was shown that a variation in the thermal conductivity of the ceramic layer, on the order of that reported in the literature for plasma sprayed ZrO2-Y2O3 coatings, can result in temperature differences in the substrate greater than 100 C, a much greater effect than that due to the presence of a bond coat or radiation from the ceramic layer. The thermal model was also used to predict the thermal response of a coated rod in order to quantify the difference in the metal temperatures between the two substrate geometries in order to explain the previously-observed increased life of coatings on rods over that on tubes. It was shown that for the short duration testing in the rocket engine, the temperature in a tube could exceed that in a rod by more than 100 C. Lastly, a two-dimensional model was developed to evaluate the effect of tangential heat transfer around the tube and its impact on reducing the stagnation point temperature. It was also shown that tangential heat transfer does not significantly reduce the stagnation point temperature, thus allowing application of a simpler, one-dimensional model for comparing measured and predicted stagnation point temperatures.

Nesbitt, James A.↗
NASA NTRS1996

Large-eddy simulation of a backward facing step flow using a least-squares spectral element method

We report preliminary results obtained from the large eddy simulation of a backward facing step at a Reynolds number of 5100. The numerical platform is based on a high order Legendre spectral element spatial discretization and a least squares time integration scheme. A non-reflective outflow boundary condition is in place to minimize the effect of downstream influence. Smagorinsky model with Van Driest near wall damping is used for sub-grid scale modeling. Comparisons of mean velocity profiles and wall pressure show good agreement with benchmark data. More studies are needed to evaluate the sensitivity of this method on numerical parameters before it is applied to complex engineering problems.

Chan, Daniel C.↗
NASA NTRS1992

Comparison of a two-dimensional adaptive-wall technique with analytical wall interference correction techniques

A two dimensional airfoil model was tested in the adaptive wall test section of the NASA Langley 0.3 meter Transonic Cryogenic Tunnel (TCT) and in the ventilated test section of the National Aeronautical Establishment Two Dimensional High Reynold Number Facility (HRNF). The primary goal of the tests was to compare different techniques (adaptive test section walls and classical, analytical corrections) to account for wall interference. Tests were conducted over a Mach number range from 0.3 to 0.8 at chord Reynolds numbers of 10 x 10(exp 6), 15 x 10(exp 6), and 20 x 10(exp 6). The angle of attack was varied from about 12 degrees up to stall. Movement of the top and bottom test section walls was used to account for the wall interference in the HRNF tests. The test results are in good agreement.

Mineck, Raymond E.↗
NASA NTRS1989

Euler and Navier-Stokes solutions for hypersonic flows

An upwind finite-element technique that uses cell-centered quantities and implicit and/or explicit time marching has been developed for computing hypersonic laminar viscous flows using adaptive unstructured grids in two and three dimensions. A perfect gas model as well as an equilibrium air model is implemented for solving high-speed flows. A first-order basic scheme and a higher-order flux-corrected transport (FCT) scheme have been implemented. This technique has been used to predict 'Type III and IV' shock interactions on a cylinder in two dimensions and a swept cylinder in three dimensions, with a view to determine the pressure and heating rate augmentation caused by an impinging shock on the leading edge of a cowl lip of an engine inlet. The predictions of wall pressure and heating rates compare very well with experimental data. The flow features are very distinctly captured with a sequence of adaptively-generated grids. Three-dimensional corner flow, typically encountered in engine inlets due to compression of the flow by ramps in the walls, is also modeled. This procedure is the first step in developing an integrated fluid, thermal, structural analysis capability for hypersonic flight vehicles like the National Aero-Space Plane.

Thareja, Rajiv R.↗
NASA NTRS1989

Radiative heat transfer in rocket thrust chambers and nozzles

Numerical models based on the discrete exchange factor (DEF) and the zonal methods for radiative analysis of rocket engines containing a radiatively participating medium have been developed. These models implement a new technique for calculating the direct exchange factors to account for possible blockage by the nozzle throat. Given the gas and surface temperature distributions, engine geometry, and radiative properties, the models compute the wall radiative heat fluxes at different axial positions. The results of sample calculations for a typical rocket engine (engine 700 at NASA), which uses RP-1 (a kerosene-type propellant), are presented for a wide range of surface and gas properties. It is found that the heat transfer by radiation can reach up to 50 percent of that due to convection. The maximum radiative heat flux is at the inner side of the engine, where the gas temperature is the highest. While the results of both models are in excellent agreement, the computation time of the DEF method is found to be much smaller.

Hammad, K. J.↗
NASA NTRS1993

Numerical computation of aerodynamics and heat transfer in a turbine cascade and a turn-around duct using advanced turbulence models

The objective of this research is to develop turbulence models to predict the flow and heat transfer fields dominated by the curvature effect such as those encountered in turbine cascades and turn-around ducts. A Navier-Stokes code has been developed using an explicit Runge-Kutta method with a two layer k-epsilon/ARSM (Algebraic Reynolds Stress Model), Chien's Low Reynolds Number (LRN) k-epsilon model and Coakley's LRN q-omega model. The near wall pressure strain correlation term was included in the ARSM. The formulation is applied to Favre-averaged N-S equations and no thin-layer approximations are made in either the mean flow or turbulence transport equations. Anisotropic scaling of artificial dissipation terms was used. Locally variable timestep was also used to improve convergence. Detailed comparisons were made between computations and data measured in a turbine cascade by Arts et al. at Von Karman Institute. The surface pressure distributions and wake profiles were predicted well by all the models. The blade heat transfer is predicted well by k-epsilon/ARSM model, as well as the k-epsilon model. It's found that the onset of boundary layer transition on both surfaces is highly dependent upon the level of local freestream turbulence intensity, which is strongly influenced by the streamline curvature. Detailed computation of the flow in the turn around duct has been carried out and validated against the data by Monson as well as Sandborn. The computed results at various streamwise locations both on the concave and convex sides are compared with flow and turbulence data including the separation zone on the inner well. The k-epsilon/ARSM model yielded relatively better results than the two-equation turbulence models. A detailed assessment of the turbulence models has been made with regard to their applicability to curved flows.

Lakshminarayana, B.↗
NASA NTRS2018

Mid-Fidelity Computational Fluid Dynamics Analysis of the Elytron 4S UAV Concept

The Elytron 4S Unmanned Aerial Vehicle (UAV) Concept was developed to combine the advantages of fixed- and rotary-wing technology. The 4S Concept is a box-wing configuration with rotors mounted on a centrally located tiltwing. The UAV is intended to be capable of both conventional takeoff and landing (CTOL) and vertical takeoff and landing (VTOL), and is envisioned to excel in UAV performance because of the combined efficiency of fixed-wing aircraft and the hover and VTOL capabilities of regular drones or quadcopters. A mid-fidelity Unsteady Reynolds-Averaged Navier-Stokes (URANS) approach using Rotorcraft CFD (RotCFD) is performed to analyze and characterize the performance of the aircraft. The flow field is coupled with a rotor model based on blade-element momentum theory to model the 4S UAV rotors. Turbulence is modeled using a realizable k- turbulence model with special wall function. The code is used to generate aerodynamic forces and moments on the body at cruise conditions, and during VTOL. The results and their uncertainties are characterized, and an angle- of-attack and sideslip sweep are computed, both with and without rotors on. Simulations are compared with the wind tunnel tests in the 7- by 10-Foot U.S. Army Wind Tunnel at NASA Ames Research Center, performed in 2017. Results show promising comparison with experimental data, despite a late change in rotor size and rudder size of the physical model that cause the expected deviations from the simulation. A slight change in the net thrust value, when rotors are modeled, is observed because of the rotor diameter increase on the physical model. A noticeable difference in the directional stability was observed because of the increased rudder surface and added strakes. These changes were implemented to improve on the design as simulated, which is observed in the results. The simulation results paved the way to the first successful flight of the UAV Concept.

Koning, Witold J. K.↗
NASA NTRS1993

Direct simulation of compressible wall-bounded turbulence

When analyzing many turbulent flows, the effects of compressibility can be neglected. Even some relatively high-speed flows, such as boundary layers generated by a supersonic aircraft, produce turbulent statistics that are similar to those found for the incompressible case. There are other situations, however, in which the non-zero divergence of the turbulence leads to behavior that is fundamentally different from that found at constant density. Examples include flows created by internal combustion engines, hypersonic flight, and supersonic combustion. It is with instances such as these that this project is concerned. In particular, we are interested in the effects of compressibility on turbulence near a smooth solid constant temperature surface; our primary objective is an increased physical understanding that can be used to improve turbulence models of wall-bounded compressible flows. With this in mind, we have begun a direct numerical simulation (DNS) study of turbulence in a plane channel. Because all of the relevant spatial and temporal scales are to be resolved, the simulations require no subgrid scale parameterization. The DNS code developed by Buell to study compressible plane Couette flow was modified to solve the compressible Navier-Stokes equations in the plane channel. The channel was chosen over the Couette flow for two reasons: (1) to avoid using the very large streamwise domains needed to adequately capture the large Couette vortical structures; and (2) to make use of previous experience by considering the compressible version of a well established case and isolate finite Mach number effects by comparing to the incompressible channel.

Coleman, Gary N.↗
NASA NTRS1991

On the application of algebraic turbulence models to high Mach number flows

In this paper the implementation of embedded-function algorithms for conventional algebraic turbulence models is described; using this methodology it is shown that the entire wall layer may be removed from the calculation algorithm (as well as near-wall mixing-length models) with no degradation in the accuracy of predicted results for skin friction. Various difficulties that are associated with the application of conventional algebraic models in high-speed flows are discussed. Alternative algebraic models are proposed which are cast in terms of the Howarth-Dorodnitsyn variable and which lead to a simple 'law of the wall' for both velocity and total enthalpy.

Degani, A. T.↗
NASA NTRS1992

Numerical analysis of hypersonic turbulent film cooling flows

As a building block, numerical capabilities for predicting heat flux and turbulent flowfields of hypersonic vehicles require extensive model validations. Computational procedures for calculating turbulent flows and heat fluxes for supersonic film cooling with parallel slot injections are described in this study. Two injectant mass flow rates with matched and unmatched pressure conditions using the database of Holden et al. (1990) are considered. To avoid uncertainties associated with the boundary conditions in testing turbulence models, detailed three-dimensional flowfields of the injection nozzle were calculated. Two computational fluid dynamics codes, GASP and FDNS, with the algebraic Baldwin-Lomax and k-epsilon models with compressibility corrections were used. It was found that the B-L model which resolves near-wall viscous sublayer is very sensitive to the inlet boundary conditions at the nozzle exit face. The k-epsilon models with improved wall functions are less sensitive to the inlet boundary conditions. The testings show that compressibility corrections are necessary for the k-epsilon model to realistically predict the heat fluxes of the hypersonic film cooling problems.

Chen, Y. S.↗
NASA NTRS1975

The self streamlining wind tunnel

A two dimensional test section in a low speed wind tunnel capable of producing flow conditions free from wall interference is presented. Flexible top and bottom walls, and rigid sidewalls from which models were mounted spanning the tunnel are shown. All walls were unperforated, and the flexible walls were positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplied the information required in the streamlining process, when run with the model present. Measurements taken at the flexible walls were used by the tunnels computer check wall contours. Suitable adjustments based on streamlining criteria were then suggested by the computer. The streamlining criterion adopted when generating infinite flowfield conditions was a matching of static pressures in the test section at a wall with pressures computed for an imaginary inviscid flowfield passing over the outside of the same wall. Aerodynamic data taken on a cylindrical model operating under high blockage conditions are presented to illustrate the operation of the tunnel in its various modes.

Goodyer, M. J.↗
NASA NTRS1987

Adaptive wall testing sections (AWTS)

The lecture starts with conventional techniques of minimizing wall interference and explains the principle of wall streamlining. The history of AWTS development is highlighted, along with the benefits of wall streamlining, including minimized boundary interference, increased model size, reduced tunnel drive power, noise, and volume, as well as multiple flow field simulations to be performed using one test section. AWTS-associated problems coming from the need to adjust the test-section boundaries for each test condition are assessed, along with the requirements of a boundary-adjustment strategy. Examples of two- and three-dimensional test sections are presented, and attention is focused on residual interference and the effects of compressibility and model lift on flexible-wall contours.

Wolf, Stephen W. D.↗
NASA NTRS1990

Navier-Stokes computation of compressible turbulent flows with a second order closure, part 1

A second order closure turbulence model for compressible flows is developed and implemented in a 2D Reynolds-averaged Navier-Stokes solver. From the beginning where a kappa-epsilon turbulence model was implemented in the bidiagonal implicit method of MACCORMACK (referred to as the MAC3 code) to the final stage of implementing a full second order closure in the efficient line Gauss-Seidel algorithm, numerous work was done, individually and collectively. Besides the collaboration itself, the final product of this work is a second order closure derived from the Launder, Reece, and Rodi model to account for near wall effects, which has been called FRAME model, which stands for FRench-AMerican-Effort. During the reporting period, two different problems were worked out. The first was to provide Ames researchers with a reliable compressible boundary layer code including a wide collection of turbulence models for quick testing of new terms, both in two equations and in second order closure (LRR and FRAME). The second topic was to complete the implementation of the FRAME model in the MAC5 code. The work related to these two different contributions is reported. dilatation in presence of stron shocks. This work, which has been conducted during a work at the Center for Turbulence Research with Zeman aimed also to cros-check earlier assumptions by Rubesin and Vandromme.

Haminh, Hieu↗