Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Boundary layer flow”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,333 records · Page 74
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Drag Reduction Devices for Aircraft (Latest Citations from the Aerospace Database)

The bibliography contains citations concerning the modeling, application, testing, and development of drag reduction devices for aircraft. Slots, flaps, fences, large-eddy breakup (LEBU) devices, vortex generators and turbines, Helmholtz resonators, and winglets are among the devices discussed. Contour shaping to ensure laminar flow, control boundary layer transition, or minimize turbulence is also covered. Applications include the wings, nacelles, fuselage, empennage, and externals of aircraft designed for high-lift, subsonic, or supersonic operation. The design, testing, and development of directional grooves, commonly called riblets, are covered in a separate bibliography.(Contains 50-250 citations and includes a subject term index and title list.)

AIRCRAFT DESIGN↗
NASA NTRS1951

Aerodynamic Characteristics of Bodies at Supersonic Speeds: A Collection of Three Papers

The three papers collected here are: 'The Effect of Nose Shape on the Drag of Bodies of Revolution at Zero Angle of Attack.', 'Base Pressure on Wings and Bodies with Turbulent Boundary Layers', and 'Flow over Inclined Bodies'. The subject of the first paper is the drag of the nose section of bodies of revolution at zero angle of attack. The main object of the second paper is to summarize the prinicpal results of the many wind tunnel and free flight measurements of base pressure on both bodies of revolution and blunt trailing edge airfoils.

Source record↗
NASA NTRS2006

Notes on Earth Atmospheric Entry for Mars Sample Return Missions

The entry of sample return vehicles (SRVs) into the Earth's atmosphere is the subject of this document. The Earth entry environment for vehicles, or capsules, returning from the planet Mars is discussed along with the subjects of dynamics, aerodynamics, and heat transfer. The material presented is intended for engineers and scientists who do not have strong backgrounds in aerodynamics, aerothermodynamics and flight mechanics. The document is not intended to be comprehensive and some important topics are omitted. The topics considered in this document include basic principles of physics (fluid mechanics, dynamics and heat transfer), chemistry and engineering mechanics. These subjects include: a) fluid mechanics (aerodynamics, aerothermodynamics, compressible fluids, shock waves, boundary layers, and flow regimes from subsonic to hypervelocity; b) the Earth s atmosphere and gravity; c) thermal protection system design considerations; d) heat and mass transfer (convection, radiation, and ablation); e) flight mechanics (basic rigid body dynamics and stability); and f) flight- and ground-test requirements; and g) trajectory and flow simulation methods.

Rivell, Thomas↗
NASA NTRS2012

Asymmetric Uncertainty Expression for High Gradient Aerodynamics

When the physics of the flow around an aircraft changes very abruptly either in time or space (e.g., flow separation/reattachment, boundary layer transition, unsteadiness, shocks, etc), the measurements that are performed in a simulated environment like a wind tunnel test or a computational simulation will most likely incorrectly predict the exact location of where (or when) the change in physics happens. There are many reasons for this, includ- ing the error introduced by simulating a real system at a smaller scale and at non-ideal conditions, or the error due to turbulence models in a computational simulation. The un- certainty analysis principles that have been developed and are being implemented today do not fully account for uncertainty in the knowledge of the location of abrupt physics changes or sharp gradients, leading to a potentially underestimated uncertainty in those areas. To address this problem, a new asymmetric aerodynamic uncertainty expression containing an extra term to account for a phase-uncertainty, the magnitude of which is emphasized in the high-gradient aerodynamic regions is proposed in this paper. Additionally, based on previous work, a method for dispersing aerodynamic data within asymmetric uncer- tainty bounds in a more realistic way has been developed for use within Monte Carlo-type analyses.

Pinier, Jeremy T↗
NASA NTRS2014

Observations on CFD Verification and Validation from the AIAA Drag Prediction Workshops

The authors provide observations from the AIAA Drag Prediction Workshops that have spanned over a decade and from a recent validation experiment at NASA Langley. These workshops provide an assessment of the predictive capability of forces and moments, focused on drag, for transonic transports. It is very difficult to manage the consistency of results in a workshop setting to perform verification and validation at the scientific level, but it may be sufficient to assess it at the level of practice. Observations thus far: 1) due to simplifications in the workshop test cases, wind tunnel data are not necessarily the “correct” results that CFD should match, 2) an average of core CFD data are not necessarily a better estimate of the true solution as it is merely an average of other solutions and has many coupled sources of variation, 3) outlier solutions should be investigated and understood, and 4) the DPW series does not have the systematic build up and definition on both the computational and experimental side that is required for detailed verification and validation. Several observations regarding the importance of the grid, effects of physical modeling, benefits of open forums, and guidance for validation experiments are discussed. The increased variation in results when predicting regions of flow separation and increased variation due to interaction effects, e.g., fuselage and horizontal tail, point out the need for validation data sets for these important flow phenomena. Experiences with a recent validation experiment at NASA Langley are included to provide guidance on validation experiments.

Computational Fluid Dynamics↗
NASA NTRS2014

Flight Tests of a Supersonic Natural Laminar Flow Airfoil

IR thermography was used to characterize the transition front on a S-NLF test article at chord Reynolds numbers in excess of 30 million Changes in transition due to Mach number, Reynolds number, and surface roughness were investigated - Regions of laminar flow in excess of 80% chord at chord Reynolds numbers greater than 14 million IR thermography clearly showed the transition front and other flow features such as shock waves impinging upon the surface A series of parallel oblique shocks, of yet unknown origin, were found to cause premature transition at higher Reynolds numbers. NASA has a current goal to eliminate barriers to the development of practical supersonic transport aircraft Drag reduction through the use of supersonic natural laminar flow (S-NLF) is currently being explored as a means of increasing aerodynamic efficiency - Tradeoffs work best for business jet class at M<2 Conventional high-speed designs minimize inviscid drag at the expense of viscous drag - Existence of strong spanwise pressure gradient leads to crossflow (CF) while adverse chordwise pressure gradients amplifies and Tollmien-Schlichting (TS) instabilities Aerion Corporation has patented a S-NLF wing design (US Patent No. 5322242) - Low sweep to control CF - dp/dx < 0 on both wing surfaces to stabilize TS - Thin wing with sharp leading edge to minimize wave drag increase due to reduction in sweep NASA and Aerion have partnered to study S-NLF since 1999 Series of S-NLF experiments flown on the NASA F-15B research test bed airplane Infrared (IR) thermography used to characterize transition - Non-intrusive, global, good spatial resolution - Captures significant flow features well

laminar flow↗
NASA NTRS2016

Comparison of Turbulent Thermal Diffusivity and Scalar Variance Models

In this study, several variable turbulent Prandtl number formulations are examined for boundary layers, pipe flow, and axisymmetric jets. The model formulations include simple algebraic relations between the thermal diffusivity and turbulent viscosity as well as more complex models that solve transport equations for the thermal variance and its dissipation rate. Results are compared with available data for wall heat transfer and profile measurements of mean temperature, the root-mean-square (RMS) fluctuating temperature, turbulent heat flux and turbulent Prandtl number. For wall-bounded problems, the algebraic models are found to best predict the rise in turbulent Prandtl number near the wall as well as the log-layer temperature profile, while the thermal variance models provide a good representation of the RMS temperature fluctuations. In jet flows, the algebraic models provide no benefit over a constant turbulent Prandtl number approach. Application of the thermal variance models finds that some significantly overpredict the temperature variance in the plume and most underpredict the thermal growth rate of the jet. The models yield very similar fluctuating temperature intensities in jets from straight pipes and smooth contraction nozzles, in contrast to data that indicate the latter should have noticeably higher values. For the particular low subsonic heated jet cases examined, changes in the turbulent Prandtl number had no effect on the centerline velocity decay.

computational fluid dynamics↗
NASA NTRS1971

Calculation of compressible nonadiabatic boundary layers in laminar, transitional and turbulent flow by the method of integral relations

A computer program was developed to do the calculations for two-dimensional or axisymmetric configurations from low speeds to hypersonic speeds with arbitrary streamwise pressure, temperature, and Mach number distributions. Options are provided for obtaining initial conditions either from experimental information or from a theoretical similarity solution. The transition region can be described either by an arbitrary distribution of intermittency or by a function based on Emmons' probability theory. Correlations were developed for use in estimating the parameters of the theoretical intermittency function. Correlations obtained from other sources are used for estimating the transition point. Comparisons were made between calculated and measured boundary layer quantities for laminar, transitional, and turbulent flows on flat plates, cones, cone flares, and a waisted body of revolution. Excellent agreement was obtained between the present theory and two other theories based on the method of finite differences. The intermittency required to reproduce some experimental heat transfer results in hypersonic flow was found to be quite different from the theoretical function. It is suggested that the simple probability theory of Emmons may not be valid for representing the intermittency of hypersonic transitional boundary layers and that the program could be useful as a tool for detailed study of the intermittency of the transition region.

Kuhn, G. D.↗
NASA NTRS2012

Correction of Excessive Precipitation over Steep Mountains in a General Circulation Model (GCM)

Excessive precipitation over steep and high mountains (EPSM) is a well-known problem in GCMs and regional climate models even at a resolution as high as 19km. The affected regions include the Andes, the Himalayas, Sierra Madre, New Guinea and others. This problem also shows up in some data assimilation products. Among the possible causes investigated in this study, we found that the most important one, by far, is a missing upward transport of heat out of the boundary layer due to the vertical circulations forced by the daytime subgrid-scale upslope winds, which in turn is forced by heated boundary layer on the slopes. These upslope winds are associated with large subgrid-scale topographic variance, which is found over steep mountains. Without such subgrid-scale heat ventilation, the resolvable-scale upslope flow in the boundary layer generated by surface sensible heat flux along the mountain slopes is excessive. Such an excessive resolvable-scale upslope flow in the boundary layer combined with the high moisture content in the boundary layer results in excessive moisture transport toward mountaintops, which in turn gives rise to excessive precipitation over the affected regions. We have parameterized the effects of subgrid-scale heated-slope-induced vertical circulation (SHVC) by removing heat from the boundary layer and depositing it in the layers higher up when topographic variance exceeds a critical value. Test results using NASA/Goddard's GEOS-5 GCM have shown that the EPSM problem is largely solved.

Chao, Winston C.↗
NASA NTRS1978

Building wakes in disturbed layers

The wind profiles in the wakes of buildings in a disturbed shear layer are investigated for conditions experienced during aircraft take-off and landing. Experiments were conducted in a pipe flow apparatus in which an equilibrium flow is disturbed by pipe roughness, simulating obstacles upwind of a building, with the disturbed flow made to flow past a second obstacle, simulating a building. Velocity and turbulence profiles of disturbed and equilibrium flows upstream and downstream of the obstacles are presented. It is concluded that the integrated momentum flow of a boundary layer behind an obstacle can be greater than the upstream momentum flow, if the exponent of the upstream velocity profile is sufficiently high and the boundary layer has been disturbed by upstream obstacles. The rise of the turbulence level behind the obstacle is also dependent on the profile exponent in a disturbed layer.

Logan, E.↗
NASA NTRS1997

Bleed Hole Flow Phenomena Studied

Boundary-layer bleed is an invaluable tool for controlling the airflow in supersonic aircraft engine inlets. Incoming air is decelerated to subsonic speeds prior to entering the compressor via a series of oblique shocks. The low momentum flow in the boundary layer interacts with these shocks, growing in thickness and, under some conditions, leading to flow separation. To remedy this, bleed holes are strategically located to remove mass from the boundary layer, reducing its thickness and helping to maintain uniform flow to the compressor. The bleed requirements for any inlet design are unique and must be validated by extensive wind tunnel testing to optimize performance and efficiency. To accelerate this process and reduce cost, researchers at the NASA Lewis Research Center initiated an experimental program to study the flow phenomena associated with bleed holes. Knowledge of these flow properties will be incorporated into computational fluid dynamics (CFD) models that will aid engine inlet designers in optimizing bleed configurations before any hardware is fabricated. This ongoing investigation is currently examining two hole geometries, 90 and 20 (both with 5-mm diameters), and various flow features.

Source record↗
NASA NTRS2011

Boundary Layer Transition over Blunt Hypersonic Vehicles Including Effects of Ablation-Induced Out-Gassing

Computations are performed to study the boundary layer instability mechanisms pertaining to hypersonic flow over blunt capsules. For capsules with ablative heat shields, transition may be influenced both by out-gassing associated with surface pyrolysis and the resulting modification of surface geometry including the formation of micro-roughness. To isolate the effects of out-gassing, this paper examines the stability of canonical boundary layer flows over a smooth surface in the presence of gas injection into the boundary layer. For a slender cone, the effects of out-gassing on the predominantly second mode instability are found to be stabilizing. In contrast, for a blunt capsule flow dominated by first mode instability, out-gassing is shown to be destabilizing. Analogous destabilizing effects of outgassing are also noted for both stationary and traveling modes of crossflow instability over a blunt sphere-cone configuration at angle of attack.

Li, Fei↗
NASA NTRS1987

Experimental study of three shock wave/turbulent boundary layer interactions

The paper presents a systematic study of the supersonic flow of a turbulent boundary layer over several compression-corner models. The wind tunnel and the compression-corner models (ramps fitted with aerodynamic fences to minimize three-dimensional effects) were identical with those used by Settles et al. (1979); constant-temperature hot-wire anemometry was used for the mass-flow measurements. The turning angles used for the compression corners were 8, 16, and 20 deg. In all three flow cases, the shock wave/turbulent flow interaction did amplify the turbulent stresses dramatically, with amplification increasing with increasing turning angle. However, different stress components were amplified by different amounts.

Smits, Alexander J.↗
NASA NTRS1989

Airfoil self-noise and prediction

A prediction method is developed for the self-generated noise of an airfoil blade encountering smooth flow. The prediction methods for the individual self-noise mechanisms are semiempirical and are based on previous theoretical studies and data obtained from tests of two- and three-dimensional airfoil blade sections. The self-noise mechanisms are due to specific boundary-layer phenomena, that is, the boundary-layer turbulence passing the trailing edge, separated-boundary-layer and stalled flow over an airfoil, vortex shedding due to laminar boundary layer instabilities, vortex shedding from blunt trailing edges, and the turbulent vortex flow existing near the tip of lifting blades. The predictions are compared successfully with published data from three self-noise studies of different airfoil shapes. An application of the prediction method is reported for a large scale-model helicopter rotor, and the predictions compared well with experimental broadband noise measurements. A computer code of the method is given.

Brooks, Thomas F.↗
NASA NTRS1984

Propagation of propeller tone noise through a fuselage boundary layer

In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Hanson, D. B.↗
NASA NTRS1977

An approximate calculation of the strong interaction on a transonic airfoil

This paper describes an interactive computational scheme that calculates the transonic flow field over an airfoil. The inviscid flow, the turbulent boundary layer in the weak interaction region, and the elliptic flow in the strong interaction region are computed separately, and the solutions are matched iteratively in a direct, second-order interaction scheme. The pressure distribution, obtained from a transonic, small-perturbation solution, is prescribed on the boundary layer and the Navier-Stokes equations and the ensuing displacement thicknesses are used to modify the inviscid flow. The best overall results are obtained by using a turbulence model that eliminates the eddy viscosity in the strong interaction region. Computed pressure distributions are compared with experimental data and with Navier-Stokes solutions. In weak interactions, the agreement with experimental data is very good, as it is also in strong interactions except for the pressures in the separated-flow region. There the agreement is qualitative only, and the pressure recovery is too high. The results in both cases are superior or equal to those of the Navier-Stokes solutions, and the present method is significantly faster.

Seginer, A.↗