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The viscous flow field near the surface of a hovering rotor blade was studied for blade twist distributions typical of a till rotor blade and a conventional helicopter rotor blade. Three blade geometries were studied, including a tilt rotor blade twist distribution (baseline), conventional helicopter rotor blade twist distribution, and the baseline twist distribution with 2 deg of precone. The results give insight into the delayed stall phenomenon often observed for highly twisted rotors. Calculations were performed for a high thrust condition near stall using the thin-layer Navier-Stokes CFD code TURNS. Effects of built-in twist on section force coefficients, skin friction, velocities, surface pressures, and boundary layer shape factor are discussed. Although the rotor thrust coefficient was nominally the same for the cases using the two twist distributions, large differences were found in the section in-plane and normal force coefficients. These preliminary results imply that the blade outboard region, rather than the inboard region, provides the majority of the performance advantage of the baseline case over the low twist case. Skin friction, velocities near the blade, and surface pressures for the two twist distributions reveal significant differences in the blade outboard region.
An approximate method for the development of flow and thermal boundary layers in the laminar region on cylinders with arbitrary cross section and transpiration-cooled walls is obtained by the use of Karman's integrated momentum equation and an analogous heat-flow equation. Incompressible flow with constant property values throughout the boundary layer is assumed. The velocity and temperature profiles within the boundary layer are approximated by expressions composed of trigonometric functions. Shape parameters for these profiles and functions necessary for the solution of the boundary-layer equations are presented as graphs so that the calculation for any specific case is reduced to the solution of two first-order differential equations. This method is applied to determine local heat-transfer coefficients and surface temperatures in the laminar flow region of the transpiration-cooled turbine blades for a given coolant flow rate, or to calculate the coolant flow distribution which is necessary in order to keep the blade temperature uniform along the surface.
Following completion of the 9- by 15-Foot Low Speed Wind Tunnel acoustic improvement modification project, a characterization of the flow field throughout the 9- by 15-foot test section was conducted in 2019. The acoustic improvement modifications project included several large-scale tunnel loop modifications and reconstruction of the 9- by 15-Foot Low Speed Wind Tunnel test section and diffuser to decrease the noise floor in the test section. Data were collected at three cross-sectional planes in the 9- by 15-ft test section using a 15-foot traversing calibration rake and a set of eight boundary layer rakes. The data from the characterization test were used to generate calibration relationships for facility operation and understand the flow quality in the new test section, including uniformity, flow angularity, turbulence intensity, and boundary layer characteristics. Additionally, the subsonic quick check rake was used to baseline the check calibration data set for the new test section.
Following completion of the 9- by 15-Foot Low Speed Wind Tunnel acoustic improvement modification project, a characterization of the flow field throughout the 9- by 15-foot test section was conducted in 2019. The acoustic improvement modifications project included several large-scale tunnel loop modifications and reconstruction of the 9- by 15-Foot Low Speed Wind Tunnel test section and diffuser to decrease the noise floor in the test section. Data were collected at three cross-sectional planes in the 9- by 15-ft test section using a 15-foot traversing calibration rake and a set of eight boundary layer rakes. The data from the characterization test were used to generate calibration relationships for facility operation and understand the flow quality in the new test section, including uniformity, flow angularity, turbulence intensity, and boundary layer characteristics. Additionally, the subsonic quick check rake was used to baseline the check calibration data set for the new test section.
Flight experiments were conducted on a single-engine turboprop aircraft fitted with a 92-in-chord, 3-ft-span natural laminar flow glove at glove section lift coefficients from 0.15 to 1.10. The boundary-layer transition measurement methods used included sublimating chemicals and surface hot-film sensors. Transition occurred downstream of the minimum pressure point. Hot-film sensors provided a well-defined indication of laminar, laminar-separation, transitional, and turbulent boundary layers. Theoretical calculations of the boundary-layer parameters provided close agreement between the predicted laminar-separation point and the measured transition location. Tollmien-Schlichting (T-S) wave growth n-factors between 15 and 17 were calculated at the predicted point of laminar separation. These results suggest that for many practical airplane cruise conditions, laminar separation (as opposed to T-S instability) is the major cause of transition in predominantly two-dimensional flows.
The present investigation is concerned with the response of a well-defined, steady, flatplate turbulent boundary layer subjected subsequently to two types of oscillations of the free-stream velocity in the test section. It is found that the mean velocity profile in the boundary layer is unaffected by imposed free-stream oscillations in the range of frequencies employed. It is the same as the one measured with a free-stream velocity distribution held steady at its mean value. The behavior of the mean velocity field is a consequence of two observations. The time-averaged Reynolds stress distribution across the boundry layer is unaffected by the imposed oscillations. The Reynolds stresses arising from the organized velocity fluctuations under imposed oscillatory conditions are negligible compared to Reynolds stress due to the random fluctuations.
The design and operational characteristics of a new test leg for the High Reynolds Number Facility at Ames Research Center are presented, and the unique features of a test section for obtaining two-dimensional airfoil data are reviewed. The new facility operates at unit Reynolds numbers ranging from 1,000,000/ft to 30,000,000/ft. Boundary-layer suction panels are used in the test section (16 in. wide by 24 in. high) to minimize side-wall interference effects. Flexible, easily adjustable upper and lower walls allow test-channel area-ruling so as to nullify Mach-number changes induced by mass removal, to correct for longitudinal boundary-layer growth, and to provide contouring compatible with the streamlines of the model in free air.
Three possible Cretaceous/Tertiary (K/T) boundary sections in the Indian subcontinent were studied for their geochemical and fossil characteristics. These include two marine sections of Meghalaya and Zanskar and one continental section of Nagpur. The Um Sohryngkew river section of Meghalaya shows a high iridium, osmium, iron, cobalt, nickel and chromium concentration in a 1.5 cm thick limonitic layer about 30 cm below the planktonic Cretaceous-Palaeocene boundary identified by the characteristic fossils. The Bottaccione and Contessa sections at Gubbio were also analyzed for these elements. The geochemical pattern at the boundary at the Um Sohryngkew river and Gubbio sections are similar but the peak concentrations and the enrichment factors are different. The biological boundary is not as sharp as the geochemical boundary and the extinction appears to be a prolonged process. The Zanskar section shows, in general, similar concentration of the siderophile, lithophile and rare earth elements but no evidence of enrichment of siderophiles has so far been observed. The Takli section is a shallow inter-trappean deposit within the Deccan province, sandwiched between flow 1 and flow 2. The geochemical stratigraphy of the inter-trappeans is presented. The various horizons of ash, clay and marl show concentration of Fe and Co, generally lower than the adjacent basalts. Two horizons of slight enrichment of iridium are found within the ash layers, one near the contact of flow 1 and other near the contact of flow 2, where iridium occurs at 170 and 260 pg/g. These levels are lower by a factor of 30 compared to Ir concentration in the K/T boundary in Meghalaya section. If the enhanced level of some elements in a few horizons of the ash layer are considered as volcanic contribution by some fractionation processes than the only elements for which it occurs are REE, Ir and possibly Cr.
Results of an isolated inlet test for NASA's GTX air-breathing launch vehicle concept are presented. The GTX is a Vertical Take-off/ Horizontal Landing reusable single-stage-to-orbit system powered by a rocket-based combined-cycle propulsion system. Tests were conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel during two entries in October 1998 and February 1999. Tests were run from Mach 2.8 to 6. Integrated performance parameters and static pressure distributions are reported. The maximum contraction ratios achieved in the tests were lower than predicted by axisymmetric Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD). At Mach 6, the maximum contraction ratio was roughly one-half of the CFD value of 16. The addition of either boundary-layer trip strips or vortex generators had a negligible effect on the maximum contraction ratio. A shock boundary-layer interaction was also evident on the end-walls that terminate the annular flowpath cross section. Cut-back end-walls, designed to reduce the boundary-layer growth upstream of the shock and minimize the interaction, also had negligible effect on the maximum contraction ratio. Both the excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls were identified as possible reasons for the discrepancy between the CFD predictions and the experiment. It is recommended that the centerbody spike and throat angles be reduced in order to lessen the induced pressure rise. The addition of a step on the cowl surface, and planar end-walls more closely approximating a plane of symmetry are also recommended. Provisions for end-wall boundary-layer bleed should be incorporated.
A small scale Mach 3.5 wind tunnel incorporating certain novel design features and intended for boundary-layer-transition research has been tested. The free stream noise intensities and spectral distributions were determined throughout the test section for several values of unit Reynolds number and for nozzle boundary layer bleed on and off. The boundary layer transition location on a slender cone and the response of this to changes in the noise environment were determined. Root mean square free stream noise levels ranged from less than one tenth up to values approaching those for conventional nozzles, with the lowest values prevailing at upstream locations within the nozzle. For low noise conditions, cone transition Reynolds numbers were in the range of those for free flight; whereas for high noise conditions, they were in the range of those in conventional tunnels.
We perform Direct Numerical Simulations (DNS) to study the acoustic freestream disturbances radiating from the turbulent boundary layers along the contoured nozzle walls of a hypersonic wind tunnel with a rectangular test section. To begin with, the effects of the spanwise end walls are suppressed by confining the spanwise computational domain to a finite segment of the overall nozzle cross section and by imposing periodic boundary conditions across that spanwise domain. Besides providing a building-block configuration to reveal par- tial effects of the enclosed acoustic environment within the wind tunnel, these computations serve as a stepping stone toward the goal of fully-3D computations including the nozzle end walls. Building upon the earlier simulations of Deegan et al. (2018), we show that the computed acoustic characteristics in the spanwise periodic simulations are insensitive to changes in the grid resolution parameters (e.g., x+ and z+). This is relative to the previous simulations that involved a coarser resolution in the streamwise and wall-normal directions, especially up- stream of the test section. Furthermore, we outline a comparison of the pressure fluctuations induced by the turbulent boundary layers over the contoured nozzle walls and several other calculations involving boundary layers over a single flat plate at nearly the same value of the edge Mach number. We also show the impact of having only one turbulent boundary layer instead of two within the computational setup. Various flow statistics, including the first and second moments of the unsteady flow field, are computed for the different flow configurations. Good comparisons of the statistics of the nozzle-wall boundary-layer turbulence and of the freestream acoustic disturbances between the simulations of coarser and finer grids confirms DNS procedure, but also the insensitivity of noise characteristics t section to the inflow turbulence generation technique.
The test section walls of the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel are known to move under thermal and pressure loads. Videogrammetry was used to measure wall motion during the summer of 2002. In addition, a laser distancemeter was used to measure the relative distance between the test section walls at a single point. Distancemeter and videogrammetry results were consistent. Data were analyzed as a function of temperature and pressure to determine their effects on wall motion. Data were collected between 50 and 100 F, 0 and 0.315 Mach, and dynamic pressures of 0 and 120 psf. The overall motion of each wall was found to be less than 0.25 in. and less than facility personnel anticipated. The results show how motion depends on the temperature and pressure inside the test section as well is the position of the boundary layer vane. The repeatability of the measurements was +/-0.06 in. This report describes the methods used to record the motion of the test section walls and the results of the data analysis. Future facility plans include the development of a suitable wall restraint system and the determination of the effects of the wall motion on tunnel calibration.
Measurements have been made of the heat transfer through a turbulent boundary layer on a convexly curved isothermal wall and on a flat plate following the curved section. Data were taken for one free-stream velocity and two different ratios of boundary layer thickness to radius of curvature delta/R = 0.051 and delta/R = 0.077. Only small differences were observed in the distribution of heat transfer rates for the two boundary layer thicknesses tested, although differences were noted in the temperature distributions within the boundary layer
In the present investigation of the application of convex-curvature viscous drag reduction to an axisymmetric body, the qualification of a modified mixing-length formulation in several complex two-dimensional and axisymmetric turbulent boundary layers that are primarily curved is followed by the design of a low-drag nose-body for low drag at high subsonic speeds. The longitudinal distribution of the cross-sectional area ratio is found to be critical to boundary layer separation, requiring implementation not over a single long span of curvature but over three short spans. The possibility of a 'net' drag reduction is examined in relation to 'equivalent' nose-bodies.
Empirical methods have been developed to estimate the test section noise levels and the outside noise radiation patterns of blowdown wind tunnels. Included are considerations of noise generation by control valves, burners, turbulent boundary layers, and exhaust jets as appropriate. Sample test section and radiation field noise estimates are presented. The external estimates are noted to be in good agreement with the limited amount of available measurements.
Measurements of mean velocity, turbulence intensity, and wall heat transfer were obtained in a 13 deg total angle of divergence conical diffuser coupled to a constant diameter recovery section. The results indicated that the boundary layer was in a nearly separated state. Turbulence intensity levels approaching 0.4 were observed in the latter stages of diffusion. The convective heat transfer was always equal to or higher than corresponding values for fully developed pipe flow at the same Reynolds number. The augmentation in heat transfer was greatest during the latter stages of diffusion where the Stanton number was nearly three times the pipe flow value.
Results of experimental studies on the effect of various factors on the transition of a supersonic boundary layer are discussed. It is shown that in supersonic wind tunnels, a significant effect on the transition of the boundary layer on a model is exerted by the scale of acoustic perturbations, which is proportional to the boundary layer displacement thickness of the working section. Experimental data obtained over a wide range of variation of flow parameters in aerodynamically similar test installations with different dimensions of the working section are generalized by means of a correlation parameter based on the displacement thickness.