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At least 1,261 records · Page 70
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Characterization of Aerodynamic Performance of Boundary-Layer-Ingesting Inlet Under Crosswind

NASA has been studying future transport concepts, envisioned to be technically realizable in the timeframe of 2020-2030, to meet environmental and performance goals. One concept receiving considerable interest involves a propulsion system embedded into a hybrid wing-body aircraft. While offering significant advantages in fuel savings and noise reduction by this concept, there are several technical challenges that are not encountered in the current fleet and must be overcome so as to deliver target performance and operability. One of these challenges is associated with an inlet system that ingests a significantly thick boundary layer, developing along the wing-body surface, into a serpentine diffuser before the flow meeting fan blades. The flow is subject to considerable total pressure loss and distorted at the fan face, much more significantly than in the inlet system of conventional aircraft. In our previous studies [1, 2], we have shown that through innovative design changes on the airframe surface, it is possible to simultaneously increase total pressure recovery and decrease distortion in the flow, without resorting to conventional penalty-ridden flow control concepts, such as vortex generator or boundary layer bleeding/suction. In the current study, we are interested in understanding the following issues: how the embedded propulsion system performs under a crosswind condition by studying in detail the flow characteristics of two inlets, the baseline and another optimized previously under the cruise condition. With the insight, it is hoped that it can help in the follow-on study by devising effective strategies to minimize flow distortion arising from the integration of an embedded-engine system into an airframe to the level acceptable to the operation and fuel consumption before 2030. To achieve these demanding goals, non-conventional concepts are called for; but technology gap is too big that it requires evolutionary approach by focusing various concepts and technologies needed in the next three generations of aircraft, respectively named as N+1, N+2, and N+3. Noticeably, considerable reduction in each category of 1 is required in N+2 (relative to Boeing 777-200 and GE90 engines) and N+3 (relative to Boeing 737-800 and CFM56-7B engines). In this study, concepts for N+2 is our interest. A concept that has potential to achieve these metrics and has been under intensive study is the hybrid wing body (HWB) airframe with a tightly integrated propulsion system, see Fig. 1. The inlet is non-circular at the entrance and the entering flow, no longer uniform or free of disturbances, and is now carrying with it a boundary layer developing along the fuselage; the inlet is thus known as boundary-layer-ingesting (BLI) inlet.

Liou, Meng-Sing↗
NASA NTRS1991

Modeling turbulent boundary layers in adverse pressure gradients

Many of the turbulent layers encountered in practical flows develop in adverse pressure gradients; hence, the dynamics of the thickening and possible separation of the boundary layer has important implications for design practices. What are the key physical processes that govern how a turbulent boundary layer responds to an adverse pressure gradient, and how should these processes be modeled? Despite the ubiquity of such flows in engineering and nature, these equations remain largely unanswered. The turbulence closure models presently used to describe these flows commonly use 'wall functions' that have ad hoc corrections for the effects of pressure gradients. There is, therefore, a practical and theoretical need to examine the effects of adverse pressure gradients on wall bounded turbulent flows in order to develop models based on sound physical principle. The evolution of a turbulent boundary layer on a flat wall with an externally imposed pressure gradient is studied.

Belcher, Stephen E.↗
NASA NTRS2021

Boundary-layer Instability Measurements on a Cone at Freestream Mach 3.5

An experimental study was conducted in the NASA (National Aeronautics and Space Administration) Langley Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7◦ half-angle cone at nominal freestream conditions: Mach 3.5, total temperature of 299.8K, and unit Reynolds numbers (millions per m) of 9.89, 13.85, 21.77, and 25.73. Instability measurements were acquired under noisy-flow and quiet-flow conditions. Pitot-pressure and calibrated hot-wire measurements were obtained using a model-integrated traverse system to document the model flow field. In noisy-flow conditions, growth rates and mode shapes achieved good agreement between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N ≈ 3.9. Under quiet-flow conditions, the most unstable first-mode instabilities as predicted by LST were measured, but this mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were less-amplified, low-frequency disturbances predicted by LST, and grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process believed to occur near the branch I location of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest unit Reynolds number, implying a transition N factor of N > 8.5.

Boundary Layer↗
NASA NTRS2020

NASA Langley Aerothermodynamic Ground Tests in Support of the Boundary Layer Transition (BOLT) Flight Experiment

The Air Force Office of Scientific Research (AFOSR) has sponsored the Boundary Layer Transition (BOLT) Experiments to investigate hypersonic boundary layer transition on a low-curvature, concave surface with swept leading edges. This paper provides a review of aerothermodynamic ground test contributions by NASA Langley Research Center to the design of the BOLT flight experiment. Several test entries into the Langley Aerothermodynamics Laboratory 20-Inch Mach 6 Air Tunnel are discussed. Global surface heating distributions on subscale BOLT models were measured using either phosphor thermography or infrared thermography at a range of model attitudes and freestream Reynolds numbers. An initial test entry provided the first experimental measurements of transition on the BOLT geometry. A second test entry investigated the effects of distributed surface roughness along the swept leading edges. A third test entry provided an evaluation of the aeroheating environment on the aerodynamic fairings and a portion of the flight vehicle downstream of the BOLT experiment. These entries were intended to support the development and design of flight hardware and instrumentation for the BOLT flight experiment.

Elizabeth F Rieken↗
DOE OSTI2024

The vertical-velocity skewness in the atmospheric boundary layer without buoyancy and Coriolis effects

One of the main features of near-neutral atmospheric boundary layer (ABL) turbulence is the positive vertical velocity skewness $Sk_w$ above the roughness sublayer or the buffer region in smooth-walls. The $Sk_w$ variations are receiving renewed interest in many climate-related parameterizations of the ABL given their significance to cloud formation and to testing sub-grid schemes for Large Eddy Simulations (LES). The vertical variations of $Sk_w$ are explored here using wind tunnel and flume experiments collected above smooth, rough, and permeable-walls in the absence of buoyancy and Coriolis effects. These laboratory experiments form a necessary starting point to probe the canonical structure of $Sk_w$ as they deal with a key limiting case (i.e., near-neutral conditions). Diagnostic models based on cumulant expansions, realizability constraints, and constant mass flux approach routinely employed in the convective boundary layer as well as prognostic models based on third-order budgets are used to explain variations in $Sk_w$ for the idealized laboratory conditions. The failure of flux-gradient relations to model $Sk_w$ from the gradients of the vertical velocity variance σ$_w^2$ are explained and corrections based on models of energy transport offered. Novel links between the diagnostic and prognostic models are also featured, especially for the inertial term in the third-order budget of the vertical velocity fluctuation. The co-spectral properties of w′/σ w vs w′ 2 /σ$_w^2$ are also presented for the first time to assess the dominant scales governing $Sk_w$ in the inner and outer layers, where w′ is the fluctuating vertical velocity and σ w is the vertical velocity standard deviation.>

Boundary layer flow↗
NASA NTRS1992

Budget analysis of the boundary layer grid flights during FIFE 1987

An aircraft grid pattern was flown by the Canadian Twin Otter to map the low-level fluxes and structure over the First International Satellite Land Surface Climatology Project (ISLSCP) Field Experiment (FIFE) research area in 1987. The time dependence and horizontal advection of heat and moisture were extracted from these flights, combined with surface flux measurements and boundary layer top measurements from radiosondes, to analyze the boundary layer budget using a mixed layer model. The results confirm the suggestion of an earlier study that the boundary layer top entrainment (when parameterized using the buoyancy flux) is nearly double the value used by many modeling studies. Both surface and aircraft data have been revised, and it now appears that the direct measurements of the sensible and latent heat fluxes by the aircraft underestimated these fluxes by about 20 percent, because of filtering and undersampling of long wavelength contributions.

Betts, A. K.↗
NASA NTRS1990

Effects of forebody geometry on subsonic boundary-layer stability

As part of an effort to develop computational techniques for design of natural laminar flow fuselages, a computational study was made of the effect of forebody geometry on laminar boundary layer stability on axisymmetric body shapes. The effects of nose radius on the stability of the incompressible laminar boundary layer was computationally investigated using linear stability theory for body length Reynolds numbers representative of small and medium-sized airplanes. The steepness of the pressure gradient and the value of the minimum pressure (both functions of fineness ratio) govern the stability of laminar flow possible on an axisymmetric body at a given Reynolds number. It was found that to keep the laminar boundary layer stable for extended lengths, it is important to have a small nose radius. However, nose shapes with extremely small nose radii produce large pressure peaks at off-design angles of attack and can produce vortices which would adversely affect transition.

Dodbele, Simha S.↗
NASA NTRS1990

Boundary layer roll circulations during FIRE

The probable mechanism underlying the development of boundary layer roll circulations are studied using wind and temperature profiles measured by the National Center for Atmospheric Research (NCAR) Electra during the stratocumulus phase of the First ISCCP Regional Experiment (FIRE). The expected, or preferred, roll orientations, horizontal wavelengths, and propagation periods are determined by finding the minimum values of the dynamic and thermodynamic forcing parameters, which here are the eddy Reynolds number (Re) and moist Rayleigh number (Ra sub m). These minimum values depend on the height z sub T of the capping temperature inversion and on the values of the Fourier coefficients of the background height-dependent vector wind profile. As input to our nonlinear spectral model, descent and ascent runs by the Electra provide for initial estimates of the inversion height and the wind profiles. In the first phase of the investigation presented here, a mechanism is said to be a probable contributor to the development of roll circulations within the stratocumulus-topped boundary layer if the modeled roll orientation and wavelengths agree with their observed values. Preliminary results using the 14-coefficient model of Haack-Hirschberg (1988) are discussed for the 7 July 1987 Electra Mission 188-A (Flight 5). This mission was flown across a sharp cloud boundary that was within a LANDSAT/SPOT scene. The stratocumulus deck was relatively solid in the eastern part of the scene, while there was a rapid decrease in cloud cover to scattered cumulus clouds aligned in streets to the west. These cloud streets were oriented nearly parallel to the mean wind direction in the layer, which was approximately 340 degrees. The hypothesis that roll circulations occurred in both the relatively clear and the cloudy regions is investigated using as model input a descent profile obtained in the relatively clear air and an ascent profile obtained in the cloudy air. Initial results for the clear air case are that the pure inflection point mode is not possible and the pure thermal mode was oriented 35 degrees to the right of the mean wind direction. The origin of this unacceptably large discrepancy between the observed and modeled results will be investigated further and the conclusions reported at the next FIRE workshop.

Shirer, Hampton N.↗
NASA NTRS1983

Flat-plate drag measurements with vortex generators in turbulent boundary layer

Direct drag measurements were obtained on a flat plate with a spanwise row of vortex generators near the leading edge, to produce an array of stream wise vortices within the approaching turbulent boundary layer. The object was to explore the possibility of modifying the large scale structure of the boundary layer through embedded longitudinal vortices with a view to obtaining a reduction in wall shear. Both obstacle and vane type vortex generators were tested at free stream velocities 40 ft/sec to 130 ft/sec corresponding to plate length Reynolds no. 0.3 million to 0.8 million with a nominal boundary layer thickness of approximately 0.6 in. at the leading edge. A few vortex generator configurations were tested both on and off the plate to measure the total drag as well as the plate drag alone. The obstacle type devices reduced the plate drag, indicating that the wake momentum defect predominated even in the presence of streamwise vortices. The vane type vortex generators however always increased the plate drag.

Rao, D. M.↗
NASA NTRS1991

Bypass transition in boundary layers including curvature and favorable pressure gradient effects

Recent studies of 2-D boundary layers undergoing bypass transition were reviewed. Bypass transition is characterized by the sudden appearance of turbulent spots in boundary layer without first the regular, observable growth of disturbances predicted by linear stability theory. There are no standard criteria or parameters for defining bypass transition, but it is known to be the mode of transition when the flow is disturbed by perturbations of sufficient amplitude.

Volino, R. J.↗
NASA NTRS2018

Laser Displacement Measurements of Fan Blades in Resonance and Flutter During the Boundary Layer Ingesting Inlet and Distortion-Tolerant Fan Test

NASA's Advanced Air Transport Technology Project is investigating boundary layer ingesting propulsors for future subsonic commercial aircraft to improve aircraft efficiency, thereby reducing fuel burn. To that end, a boundary layer ingesting inlet and distortion-tolerant fan stage was designed, fabricated, and tested within the 8' x 6' Supersonic Wind Tunnel at NASA Glenn Research Center. Because of the distortion in the air flow over the fan, the blades were designed to withstand a much higher aerodynamic forcing than for a typical clean flow. The blade response for several resonance modes were measured during start-up and shutdown, as well as at near 85% design speed. Flutter in the first bending mode was also observed in the fan at the design speed, at an off-design condition, although instabilities were difficult to instigate with this fan in general. Blade vibrations were monitored through twelve laser displacement probes that were placed around the inner circumference of the casing, at the blade leading and trailing edges. These probes captured the movement of all the blades during the entire test. Results are presented for various resonance mode amplitudes, frequencies and damping, as well as flutter amplitudes and frequency. Benefits and disadvantages of laser displacement probe measurements versus strain gage measurements are discussed.

NSMS↗
NASA NTRS1976

Characteristics of the boundary-layer equations of the minimum time-to-climb problem

In many singular perturbation solutions of optimal control problems, the most difficult numerical task is to solve the boundary-layer equations. However, these equations have a special structure that may often be used to expedite their solution. This paper begins by noting the general nature of the boundary-layer equations for optimal control problems. These results are then applied to the aircraft minimum time-to-climb problem. A specific numerical example is considered to illustrate the characteristics of the solution of the boundary-layer equations for this problem.

Ardema, M. D.↗
NASA NTRS1973

A boundary layer approach to the analysis of atmospheric motion over a surface obstruction

A boundary layer approach for the solution of the flow field induced over a two-dimensional surface obstruction, such as a building or other man-made structure, is proposed. Adopting a specific geometry in the form of a semi-elliptical cylinder, the characteristics of atmospheric shear flow over a rough terrain are coupled with the well-known boundary layer equations. Two approaches are presented to incorporate the pressure field and boundary conditions which exist within the large viscous region over the obstruction. The first considers a region in the immediate vicinity of the body in which the pressure distribution and outer boundary condition on the velocity are computed from potential theory for flow over the elliptical cylinder. The second approach considers a much larger region of influence, extending from the surface to the undisturbed flow at large heights above the obstruction. Methods which appear to provide an improved theoretical model of the flow over the ellipse, such as a technique for simulating the effect of the seperation regions upstream and downstream of the body, are also presented.

Frost, W.↗
NASA NTRS1953

Comparison of secondary flows and boundary-layer accumulations in several turbine nozzles

An investigation was made of losses and secondary flows in three different turbine nozzle configurations in annular cascade. Appreciable outer shroud loss cores (passage vortices) were found to exist at the discharge of blades which had thickened suction surface boundary layers near the outer shroud. Blade designs having thinner boundary layers did not show such outer shroud loss cores, but indicated greater inward radial flow of low momentum air, in the wake loss is to this extent an indication of the presence or absence of radial flow. The blade wake was a combination of profile loss and low momentum air from the outer shroud, and the magnitude of the wake loss is to this extent an indication of the presence or absence of radial flow. At a high Mach number, shock-boundary-layer thickening on the blade suction surfaces provided an additional radial flow path for low momentum air, which resulted in large inner shroud loss regions accompanied by large deviations from design values of discharge angle. (author)

Kofskey, Milton G↗
NASA NTRS1990

Flow separation in oblique shock wave turbulent boundary layer interactions

Available experimental data on flow separation in shock wave turbulent boundary layer interactions are reviewed at hypersonic and supersonic speeds in two dimensional and axisymmetric interactions. The conditions leading to flow separation and the subsequent changes in the flow are discussed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. Pertinent experimental data for the separated flow characteristics in turbulent boundary layer oblique shock wave interactions are also presented and discussed. Empirical correlations for separated flows are also reviewed.

Hamed, A.↗
NASA NTRS1938

Determination of Boundary-Layer Transition on Three Symmetrical Airfoils in the NACA Full-Scale Wind Tunnel

For the purpose of studying the transition from laminar to turbulent flow, boundary-layer measurements were made in the NACA full-scale wind tunnel on three symmetrical airfoils of NACA 0009, 0012, and 0018 sections. The effects of variations in lift coefficient, Reynolds number, and airfoil thickness on transition were investigated. Air speed in the boundary layer was measured by total-head tubes and by hot wires; a comparison of transition as indicated by the two techniques was obtained. The results indicate no unique value of Reynolds number for the transition, whether the Reynolds number is based upon the distance along the chord or upon the thickness of the boundary layer at the transition point. In general, the transition is not abrupt and occurs in a region that varies in length as a function of the test conditions.

Silverstein, Abe↗
NASA NTRS1976

Shock wave/turbulent boundary-layer interactions with and without surface cooling

An experimental investigation was conducted to delineate the structure of the flowfield and temperature distributions in a shock wave/turbulent boundary-layer interaction with and without surface cooling. The Mach number upstream was about 3.5, and the wave angle was 23 deg. The wall to stagnation temperature ratio was 0.44 with cooling and 1.1 with heating. A detailed map of the interaction flowfields deduced from numerous boundary-layer traversing stations revealed the influence of wall cooling on the flowfield, wave structure, and size of the flow separation region. With surface cooling, the size of the separation region was much smaller, and the separation and reflected shock waves merged together near the edge of the velocity boundary layer, extending into the freestream as one wave.

Back, L. H.↗
NASA NTRS1959

Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration

Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

Maki, Ralph L.↗