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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

The turbulent thermal boundary layer with an abrupt change from a rough to a smooth wall

The work reported here was motivated by concern over the use of smooth heat flux gages for heat transfer measurements on the otherwise rough turbine blades. Stanton number distributions and boundary layer profiles of mean temperature, mean velocity, and turbulence intensity are reported for a surface with a step change from a rough to a smooth surface. In most cases, the Stanton number immediately downstream of the change in roughness drops below the all-smooth-wall data at the same x-Reynolds number. The alignment of the smooth surface between the bases and crests of the roughness elements is shown to have only a weak effect on the Stanton number distribution. It is concluded that the use of smooth heat flux gages on otherwise rough surfaces can cause large errors. It is recommended that heat transfer data collected in this manner be used with caution.

Taylor, Robert P.↗
NASA NTRS1991

Hypersonic cone flow predictions using an implicit upwind space-marching code

Solutions have been computed and results are presented for Problem 1, the case of Mach 9 transitional flow past a 7 deg half-angle cone at zero incidence. The solutions were computed using a code developed for the integration of the parabolized Navier-Stokes equations. The algorithm employed in the code is based on a Roe-type flux-difference-splitting scheme applied following a finite-volume approach. The basic algorithm has been modified to make it implicit and second-order accurate in the crossflow directions. Results are presented in terms of surface pressure and heat transfer as well as boundary layer profiles of pitot pressure, Mach number, and tangential velocity. The case was recalculated several times in an effort to determine sensitivities to such parameters as grid density, wall temperature, turbulence model parameters, as well as freestream expansion. Comparisons with the experimental data are presented and discussed.

Lawrence, Scott L.↗
NASA NTRS1938

A preliminary investigation of boundary-layer transition along a flat plate with adverse pressure gradient

Boundary-layer surveys were made throughout the transition region along a smooth flat plate placed in an airstream of practically zero turbulence and with an adverse pressure gradient. The boundary-layer Reynolds number at the laminar separation point was varied from 1,800 to 2,600. The test data, when considered in the light of certain theoretical deductions, indicated that transition probably began with separation of the laminar boundary layer. The extent of the transition region, defined as the distance from a calculated laminar separation point to the position of the first fully developed turbulent boundary-layer profile, could be expressed as a constant Reynolds number run of approximately 70,000. Some speculations are presented concerning the application of the foregoing concepts, after certain assumptions have been made, to the problem of the connection between transition on the upper surface of an airfoil at high angles of attack and the maximum lift.

Von Doenhoff, Albert E↗
NASA NTRS1954

On the three-dimensional instability of laminar boundary layers on concave walls

A study is made of the stability of laminar boundary-layer profiles on slightly curved walls relative to small disturbances that result from vortices whose axes are parallel to the principal direction of flow. The result is an eigenvalue problem by which, for a given undisturbed flow at a prescribed wall, the amplification or decay is computed for each Reynolds number and each vortex thickness. For neutral disturbances (zero amplification) a critical Reynolds number is determined for each vortex distribution. The numerical calculation produces amplified disturbances on concave walls only.

FLOW, LAMINAR↗
NASA NTRS1993

Detailed measurements in the SSME high pressure fuel turbine with smooth rotor blades

Several tests of the Rocketdyne configuration of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) Turbine have been completed in the Turbine Test Equipment (TTE) at Marshall Space Flight Center. The tests involved using scaled performance parameters and model measurements to predict the performance of the turbine. The overall performance has been the primary objective of the tests to date, but more detailed measurements are also of interest. During the most recent test of the Rocketdyne configuration of the HPFTP turbine with smooth rotor blades, several different measurement techniques were used to study the turbine inlet and exit velocity profiles, boundary layer thicknesses, turbulence intensities, etc. Data has been obtained using various hot film probes and three-hole cobra probes. Laser Velocimeter measurements were also made. The test plan and test data will be presented and discussed as well as lessons learned on how to obtain the various types of data.

Hudson, Susan T.↗
NASA NTRS1995

A parametric study of supersonic laminar flow for swept wings using linear stability analysis

A parametric study to predict the extent of laminar flow on the upper surface of a generic swept-back wing (NACA 64A010 airfoil section) at supersonic speeds was conducted. The results were obtained by using surface pressure predictions from an Euler/Navier-Stokes computational fluid dynamics code coupled with a boundary layer code, which predicts detailed boundary layer profiles, and finally with a linear stability code to determine the extent of laminar flow. The parameters addressed are Reynolds number, angle of attack, and leading-edge wing sweep. The results of this study show that an increase in angle of attack, for specific Reynolds numbers, can actually delay transition. Therefore, higher lift capability, caused by the increased angle of attack, as well as a reduction in viscous drag due to the delay in transition is possible for certain flight conditions.

Cummings, Russell M.↗
NASA NTRS1996

F-15B/Flight Test Fixture 2: A Test Bed for Flight Research

NASA Dryden Flight Research Center has developed a second-generation flight test fixture for use as a generic test bed for aerodynamic and fluid mechanics research. The Flight Test Fixture 2 (FTF-2) is a low-aspect-ratio vertical fin-like shape that is mounted on the centerline of the F-I5B lower fuselage. The fixture is designed for flight research at Mach numbers to a maximum of 2.0. The FTF-2 is a composite structure with a modular configuration and removable components for functional flexibility. This report documents the flow environment of the fixture, such as surface pressure distributions and boundary-layer profiles, throughout a matrix of conditions within the F-15B/FTF-2 flight envelope. Environmental conditions within the fixture are presented to assist in the design and testing of future avionics and instrumentation. The intent of this document is to serve as a user's guide and assist in the development of future flight experiments that use the FTF-2 as a test bed. Additional information enclosed in the appendices has been included to assist with more detailed analyses, if required.

David M Richwine↗
NASA NTRS2000

Design and Evaluation of a New Boundary-Layer Rake for Flight Testing

A new boundary-layer rake has been designed and built for flight testing on the NASA Dryden Flight Research Center F-15B/Flight Test Fixture. A feature unique to this rake is its curved body, which allows pitot tubes to be more densely clustered in the near-wall region than conventional rakes allow. This curved rake design has a complex three-dimensional shape that requires innovative solid-modeling and machining techniques. Finite-element stress analysis of the new design shows high factors of safety. The rake has passed a ground test in which random vibration measuring 12 g rms was applied for 20 min in each of the three normal directions. Aerodynamic evaluation of the rake has been conducted in the NASA Glenn Research Center 8 x 6 Supersonic Wind Tunnel at Mach 0-2. The pitot pressures from the new rake agree with conventional rake data over the range of Mach numbers tested. The boundary-layer profiles computed from the rake data have been shown to have the standard logarithmic-law profile. Skin friction values computed from the rake data using the Clauser plot method agree with the Preston tube results and the van Driest II compressible skin friction correlation to approximately +/-5 percent.

Bui, Trong T.↗
NASA NTRS2000

Pegasus(Registered trademark) Wing-Glove Experiment to Document Hypersonic Crossflow Transition: Measurement System and Selected Flight Results

In a recent flight experiment to study hypersonic crossflow transition, boundary layer characteristics were documented. A smooth steel glove was mounted on the first stage delta wing of Orbital Sciences Corporation's Pegasus (R) launch vehicle and was flown at speeds of up to Mach 8 and altitudes of up to 250,000 ft. The wing-glove experiment was flown as a secondary payload off the coast of Florida in October 1998. This paper describes the measurement system developed. Samples of the results obtained for different parts of the trajectory are included to show the characteristics and quality of the data. Thermocouples and pressure sensors (including Preston tubes, Stanton tubes, and a "probeless" pressure rake showing boundary layer profiles) measured the time-averaged flow. Surface hot-films and high-frequency pressure transducers measured flow dynamics. Because the vehicle was not recoverable, it was necessary to design a system for real-time onboard processing and transmission. Onboard processing included spectral averaging. The quality and consistency of data obtained was good and met the experiment requirements.

Bertelrud, Arild↗
NASA NTRS2001

Isolating Curvature Effects in Computing Wall-Bounded Turbulent Flows

The flow over the zero-pressure-gradient So-Mellor convex curved wall is simulated using the Navier-Stokes equations. An inviscid effective outer wall shape, undocumented in the experiment, is obtained by using an adjoint optimization method with the desired pressure distribution on the inner wall as the cost function. Using this wall shape with a Navier-Stokes method, the abilities of various turbulence models to simulate the effects of curvature without the complicating factor of streamwise pressure gradient can be evaluated. The one-equation Spalart-Allmaras turbulence model overpredicts eddy viscosity, and its boundary layer profiles are too full. A curvature-corrected version of this model improves results, which are sensitive to the choice of a particular constant. An explicit algebraic stress model does a reasonable job predicting this flow field. However, results can be slightly improved by modifying the assumption on anisotropy equilibrium in the model's derivation. The resulting curvature-corrected explicit algebraic stress model possesses no heuristic functions or additional constants. It lowers slightly the computed skin friction coefficient and the turbulent stress levels for this case (in better agreement with experiment), but the effect on computed velocity profiles is very small.

Rumsey, Christopher L.↗
NASA NTRS2001

GAP Noise Computation By The CE/SE Method

A typical gap noise problem is considered in this paper using the new space-time conservation element and solution element (CE/SE) method. Implementation of the computation is straightforward. No turbulence model, LES (large eddy simulation) or a preset boundary layer profile is used, yet the computed frequency agrees well with the experimental one.

Loh, Ching Y.↗
NASA NTRS1995

Study of the Effect of Two-Dimensional Cavities on the Boundary Layer in an Adverse Pressure Gradient

One concept for high lift aerodynamics is the trapped-vortex. Recently it has been suggested that several spoilers located at different chordwise locations on an airfoil could be deployed to form several cavities in the chordwise direction. This may provide a means of increasing upper surface camber and thereby increase wing lift. It is envisioned that the cavities would be located on an airfoil in its pressure recovery region. Investigations of the effect of one or more cavities on the adjacent boundary layer in an adverse boundary layer were not found in a literature survey. Since flow separation can be caused by flow over cavities, boundary layer measurements were made in the vicinity of cavities which were located in an adverse pressure gradient to study the associated flow in an effort to identify any benefits. The experimental data from the present investigation indicate a minimal impact on the boundary layer profiles is caused by the presence of cavities in either a constant pressure freestream or an adverse pressure gradient. The data from the experimental investigation has been obtained and analyzed. The computational investigation using an incompressible Navier-Stokes code show that reasonable results are being obtained when they are compared with the experimental data.

Margason, Richard J.↗
NASA NTRS1999

Computational Analysis of Compressibility Effects on a High-Lift Wing

The objective of this study was to investigate compressibility effects on a high-lift flowfield by simulating the flow about a three-dimensional multi-element wing. The computations were performed by solving both the incompressible and compressible Navier-Stokes equations (using the INS3D and OVERFLOW codes) on structured, overset grids. Turbulence was modeled via the one-equation, fully turbulent Spalart-Allmaras model. The computational results were validated with surface pressure measurements acquired at the NASA Ames 7- by 10-Foot Wind Tunnel. The geometry used for all computations consisted of an unswept wing in a landing configuration with a half-span flap and a three-quarter-span slat mounted inside a rectangular duct approximating the wind tunnel walls. The solutions were carefully examined to account for effects due to differences in algorithms. Compressibility effects were demonstrated by comparing surface particle traces, sectional pressure coefficient and boundary layer profile plots. It was found that small regions of compressibility near the slat and main-element leading edge can largely impact the flow. Even small compressibility regions can have significant global effects on the circulation and separation of each of the high-lift elements.

Baker, M. David↗
NASA NTRS2004

Advanced Subsonic Technology (AST) 22-Inch Low Noise Research Fan Rig Preliminary Design of ADP-Type Fan 3

This report presents results of the work completed on the preliminary design of Fan 3 of NASA s 22-inch Fan Low Noise Research project. Fan 3 was intended to build on the experience gained from Fans 1 and 2 by demonstrating noise reduction technology that surpasses 1992 levels by 6 dB. The work was performed as part of NASA s Advanced Subsonic Technology (AST) program. Work on this task was conducted in the areas of CFD code validation, acoustic prediction and validation, rotor parametric studies, and fan exit guide vane (FEGV) studies up to the time when a NASA decision was made to cancel the design, fabrication and testing phases of the work. The scope of the program changed accordingly to concentrate on two subtasks: (1) Rig data analysis and CFD code validation and (2) Fan and FEGV optimization studies. The results of the CFD code validation work showed that this tool predicts 3D flowfield features well from the blade trailing edge to about a chord downstream. The CFD tool loses accuracy as the distance from the trailing edge increases beyond a blade chord. The comparisons of noise predictions to rig test data showed that both the tone noise tool and the broadband noise tool demonstrated reasonable agreement with the data to the degree that these tools can reliably be used for design work. The section on rig airflow and inlet separation analysis describes the method used to determine total fan airflow, shows the good agreement of predicted boundary layer profiles to measured profiles, and shows separation angles of attack ranging from 29.5 to 27deg for the range of airflows tested. The results of the rotor parametric studies were significant in leading to the decision not to pursue a new rotor design for Fan 3 and resulted in recommendations to concentrate efforts on FEGV stator designs. The ensuing parametric study on FEGV designs showed the potential for 8 to 10 EPNdB noise reduction relative to the baseline.

Jeracki, Robert J.↗
NASA NTRS2006

Acoustic Characteristics of Various Treatment Panel Designs Specific to HSCT Mixer-Ejector Application

The development process of liner design methodology is described in several reports. The results of the initial effort of concept development, screening, laboratory testing of various liner concepts, and preliminary correlation (generic data) are presented in a report Acoustic Characteristics of Various Treatment Panel Designs for HSCT Ejector Liner Acoustic Technology Development Program. The second phase of laboratory test results of more practical concepts and their data correlations are presented in this report (product specific). In particular, this report contains normal incidence impedance measurements of several liner types in both a static rig and in a high temperature flow duct rig. The flow duct rig allows for temperatures up to 400 F with a grazing flow up to Mach 0.8. Measurements of impedance, DC flow resistance, and in the flow rig cases, impact of the liner on boundary layer profiles are documented. In addition to liner rig tests, a limited number of tests were made on liners installed in a mixer-Ejector nozzle to confirm the performance of the liner prediction in an installed configuration.

Salikuddin, M.↗
NASA NTRS2007

Review of Cranked-Arrow Wing Aerodynamics Project: Its International Aeronautical Community Role

This paper provides a brief history of the F-16XL-1 aircraft, its role in the High Speed Research (HSR) program and how it was morphed into the Cranked Arrow Wing Aerodynamics Project (CAWAP). Various flight, wind-tunnel and Computational Fluid Dynamics (CFD) data sets were generated during the CAWAP. These unique and open flight datasets for surface pressures, boundary-layer profiles and skinfriction distributions, along with surface flow data, are described and sample data comparisons given. This is followed by a description of how the project became internationalized to be known as Cranked Arrow Wing Aerodynamics Project International (CAWAPI) and is concluded by an introduction to the results of a 4 year CFD predictive study of data collected at flight conditions by participating researchers.

Lamar, John E.↗
NASA NTRS2007

USM3D Unstructured Grid Solutions for CAWAPI at NASA LaRC

In support the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) to improve the Technology Readiness Level of flow solvers by comparing results with measured F-16XL-1 flight data, NASA Langley employed the TetrUSS unstructured grid solver, USM3D, to obtain solutions for all seven flight conditions of interest. A newly available solver version that incorporates a number of turbulence models, including the two-equation linear and non-linear k-epsilon, was used in this study. As a first test, a choice was made to utilize only a single grid resolution with the solver for the simulation of the different flight conditions. Comparisons are presented with three turbulence models in USM3D, flight data for surface pressure, boundary-layer profiles, and skin-friction results, as well as limited predictions from other solvers. A result of these comparisons is that the USM3D solver can be used in an engineering environment to predict flow physics on a complex configuration at flight Reynolds numbers with a two-equation linear k-epsilon turbulence model.

Lamar, John E.↗
NASA NTRS2006

Advances in Raman Lidar Measurements of Water Vapor, Cirrus Clouds and Carbon Dioxide

Narrow-band interference filters with improved transmission in the ultraviolet have been developed under NASA-funded research and used in the Raman Airborne Spectroscopic Lidar (RASL) in ground- based, upward-looking tests. RASL is an airborne Raman Lidar system designed to measure water vapor mixing ratio, and aerosol backscatter/extinction/depolarization. It also possesses the capability to make experimental measurements of cloud liquid water and carbon dioxide. It is being prepared for first flight tests during the summer of 2006. With the newly developed filters installed in RASL, measurements were made of atmospheric water vapor, cirrus cloud optical properties and carbon dioxide that improve upon any previously demonstrated using Raman lidar. Daytime boundary layer profiling of water vapor mixing ratio is performed with less than 5% random error using temporal and spatial resolution of 2-minutes and 60 - 210, respectively. Daytime cirrus cloud optical depth and extinction- to-backscatter ratio measurements are made using 1-minute average. Sufficient signal strength is demonstrated to permit the simultaneous profiling of carbon dioxide and water vapor mixing ratio into the free troposphere during the nighttime. Downward-looking from an airborne RASL should possess the same measurement statistics with approximately a factor of 5 - 10 decrease in averaging time. A description of the technology improvements are provided followed by examples of the improved Raman lidar measurements.

Whiteman, David N.↗