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At least 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Boeing Vertol noise reduction experience

Boeing Vertol Model 234, CH-47D and 347 tandem rotor helicopters demonstrate noise levels that comply with the overall noise standard requirements. The helicopters achieve these noise levels because they are, by design, free of impulsive noise in level flight and climb. Avoiding an impulsive noise signature is achieved by developing and applying configuration design criteria that eliminate blade vortex intersections and recognize advancing blade tip noise limitations in high speed forward flight. Airfoil design, rotor separation and rotor tipspeed selection all contribute to the noise control. Compliance with the approach element of the standards is marginal at best. The approach problem is not appreciably alleviated by reduced tipspeed and airfoil design. Since it is a single rotor phenomenon, it is not alleviated by rotor separation in the tandem configuration. Tip shapes may provide the answer, but to date this approach has not led to a satisfactory solution.

Walls, W. W.↗
NASA NTRS1985

Development and testing of tip devices for horizontal axis wind turbines

A theoretical and field experimental program has been carried out to investigate the use of tip devices on horizontal axis wind turbine rotors. The objective was to improve performance by the reduction of tip losses. While power output can always be increased by a simple radial tip extension, such a modification also results in an increased gale load both because of the extra projected area and longer moment arm. Tip devices have the potential to increase power output without such a structural penalty. A vortex lattice computer model was used to optimize three basic tip configuration types for a 25 kW stall limited commercial wind turbine. The types were a change in tip planform, and a single-element and double-element nonplanar tip extension (winglets). A complete data acquisition system was developed which recorded three wind speed components, ambient pressure, temperature, and turbine output. The system operated unattended and could perform real-time processing of the data, displaying the measured power curve as data accumulated in either a bin sort mode or polynomial curve fit. Approximately 270 hr of perormance data were collected over a three-month period. The sampling interval was 2.4 sec; thrus over 400,000 raw data points were logged. Results for each of the three new tip devices, compared with the original tip, showed a small decrease (of the order of 1 kW) in power output over the measured range of wind speeds from cut-in at about 4 m/s to over 20 m/s, well into the stall limiting region. Changes in orientation and angle-of-attack of the winglets were not made. For aircraft wing tip devices, favorable tip shapes have been reported and it is likely that the tip devices tested in this program did not improve rotor performance because they were not optimally adjusted.

Gyatt, G. W.↗
NASA NTRS1994

Water tunnel flow visualization study of a 4.4 percent scale X-31 forebody

A water-tunnel test of a 4.4 percent-scale, forebody-only model of the X-31 aircraft with different forebody strakes and nosebooms has been performed in the Flow Visualization Facility at the NASA Dryden Flight Research Center. The focus of the study was to determine the relative effects of the different configurations on the stability and symmetry of the high-angle-of-attack forebody vortex flow field. The clean, noseboom-off configuration resisted the development of asymmetries in the primary vortices through 70 deg angle of attack. The wake of the X-31 flight test noseboom configuration significantly degraded the steadiness of the primary vortex cores and promoted asymmetries. An alternate L-shaped noseboom mounted underneath the forebody had results similar to those seen with the configuration, enabling stable, symmetrical vortices up to 70 deg angle of attack. The addition of strakes near the radome tip along the waterline increased the primary vortex strength while it simultaneously caused the vortex breakdown location co move forward. Forebody strakes did not appear to significantly reduce the asymmetries in the forebody vortex field in the presence of the flight test noseboom.

Cobleigh, Brent R.↗
NASA NTRS2016

PSF Calibration Using the Lyot-Based Low Order Wavefront Sensor Telemetry - First Simulations

One of the factor that derives the performance of small inner working angle (IWA) coronagraphs is the quality of the low-order wavefront calibration and control. The uncorrected residuals scatter the starlight and produces intensity fluctuations, which as a result creates dynamic speckle noise in the focal plane. To improve post processing of the science images, the low-order telemetry of the residuals left uncorrected by the control loop can be used to calibrate the amount of starlight leakage at small angular separations. In this proceeding, we present the preliminary simulations of point spread function calibration using the Lyot-based low-order wavefront sensor measurements (tip-tilt only) for a vector vortex coronagraph.

Low-order wavefront aberrations↗
NASA NTRS1992

Lift enhancement by trapped vortex

The viewgraphs and discussion of lift enhancement by trapped vortex are provided. Efforts are continuously being made to find simple ways to convert wings of aircraft from an efficient cruise configuration to one that develops the high lift needed during landing and takeoff. The high-lift configurations studied here consist of conventional airfoils with a trapped vortex over the upper surface. The vortex is trapped by one or two vertical fences that serve as barriers to the oncoming stream and as reflection planes for the vortex and the sink that form a separation bubble on top of the airfoil. Since the full three-dimensional unsteady flow problem over the wing of an aircraft is so complicated that it is hard to get an understanding of the principles that govern the vortex trapping process, the analysis is restricted here to the flow field illustrated in the first slide. It is assumed that the flow field between the two end plates approximates a streamwise strip of the flow over a wing. The flow between the endplates and about the airfoil consists of a spanwise vortex located between the suction orifices in the endplates. The spanwise fence or spoiler located near the nose of the airfoil serves to form a separated flow region and a shear layer. The vorticity in the shear layer is concentrated into the vortex by withdrawal of fluid at the suction orifices. As the strength of the vortex increases with time, it eventually dominates the flow in the separated region so that a shear or vertical layer is no longer shed from the tip of the fence. At that point, the vortex strength is fixed and its location is such that all of the velocity contributions at its center sum to zero thereby making it an equilibrium point for the vortex. The results of a theoretical analysis of such an idealized flow field are described.

Rossow, Vernon J.↗
NASA NTRS1983

An aeroacoustic model for high-speed, unsteady blade-vortex interaction

A three-dimensional aeroacoustic model is developed to predict the sound pulse radiated by the passage of a helicopter blade over a potential vortex. The linearized analysis assumes that either the blade-vortex separation is small or that the blade-tip Mach number is close to 1, or both, so that an acoustically noncompact situation exists. The three-dimensional blade loading due to blade-vortex interaction is constructed through a spanwise superposition of two-dimensional solutions with strength linearly increasing from hub to tip. Such a loading overestimates somewhat the strength of tip region dipoles in the acoustic calculation that follows. The final expression for the predicted far-field signature is obtained in closed form, and thus permits a relatively inexpensive calculation of the directivity of peak acoustic pressures in three dimensions.

Martinez, R.↗
NASA NTRS1974

Wind-tunnel tests and computer simulations of buoyant wing-tip vortices

It is demonstrated that wing-tip vortices with buoyant cores can be made in a wind tunnel. Wing-tip vortices generated by an aircraft often become buoyant, because of heating effects due to the jet exhaust or in connection with a descending motion in a stable atmosphere. A description of wind-tunnel studies is given, taking into account the experimental conditions, the transition to turbulence, and the drift of the buoyant vortex. A computer simulation of wing-tip vortices is also discussed.

Costen, R. C.↗
NASA NTRS1984

An experimental investigation of the chopping of helicopter main rotor tip vortices by the tail rotor

The chopping of helicopter main rotor tip vortices by the tail rotor was experimentally investigated. This is a problem of blade vortex interaction (BVI) at normal incidence where the vortex is generally parallel to the rotor axis. The experiment used a model rotor and an isolated vortex and was designed to isolate BVI noise from other types of rotor noise. Tip Mach number, radical BVI station, and free stream velocity were varied. Fluctuating blade pressures, farfield sound pressure level and directivity, velocity field of the incident vortex, and blade vortex interaction angles were measured. Blade vortex interaction was found to produce impulsive noise which radiates primarily ahead of the blade. For interaction away from the blade tip, the results demonstrate the dipole character of BVI radiation. For BVI close to the tip, three dimensional relief effect reduces the intensity of the interaction, despite larger BVI angle and higher local Mach number. Furthermore, in this case, the radiation patern is more complex due to diffraction at and pressure communication around the tip.

Ahmadi, A. R.↗
NASA NTRS1995

A study of blade-vortex interaction aeroacoustics utilizing an independently generated vortex

This paper presents results from an experimental study of rotor blade-vortex interaction (BVI) aerodynamics and acoustics. The experiment utilized an externally generated vortex interacting with a two-bladed rotor operating at zero thrust to minimize the influence of the rotor's own wake. The rotor blades were instrumented with a total of 60 absolute pressure transducers at three spanwise and ten chordwise stations on both the upper and lower surfaces. Acoustic data were obtained with fixed near-field microphones as well as a movable array of far-field microphones. The test was carried out in the acoustically treated test section of the NASA Ames 80- by 120-foot Wind Tunnel. Several parameters which influences BVI, such as vortex-rotor separation distance, vortex strength, and vortex sense (swirl direction), as well as rotor tip Mach number and advance ratio, were varied. Simultaneous measurements were obtained of blade surface pressure distributions, near-field acoustics, and far-field acoustics during the vortex-blade encounters.

Kitaplioglu, C.↗
NASA NTRS1979

Investigation of aerodynamic characteristics of subsonic wings

An analytical strake design procedure is investigated. A numerical solution to the governing strake design equation is used to generate a series of strakes which are tested in a water tunnel to study their vortex breakdown characteristics. The strakes are scaled for use on a half-scale model of the NASA-LaRC general research fuselage with a 44 degrees trapezoidal wing. An analytical solution to the governing design equation is obtained. The strake design procedure relates the potential-flow leading-edge suction and pressure distributions to vortex stability. Several suction distributions are studied and those which are more triangular and peak near the tip generate strakes that reach higher angles of attack before vortex breakdown occurs at the wing trailing edge. For the same suction distribution, a conical rather than three dimensional pressure specification results in a better strake shape as judged from its vortex breakdown characteristics.

Dejarnette, F. R.↗
NASA NTRS1987

Low-speed wind tunnel study of longitudinal stability and usable-lift improvement of a cranked wing

An exploratory low-speed investigation of a 70 deg/46 deg cranked-wing planform was undertaken to evaluate two vortex-control concepts aimed at alleviating a severe pitch up which limits the usable lift well below the C(sub L,max) of the basic wing. One concept was a strake-like extension introduced across the wing crank, whose vortex helps to stabilize the outer-wing flow and alleviate tip stall. The other was a lower-surface cavity flap employed to trap a vortex just beneath the inboard leading edge, resulting in reduced vortex lift over the inner-wing panel. Each of these concepts was shown to eliminate the high-alpha pitch up, potentially raising the maximum usable lift of the cranked wing practically to its C(sub L,max) value.

Rao, Dhanvada M.↗
NASA NTRS1992

An experimental study of heat transfer in a large-scale turbine rotor passage

An experimental study of the heat transfer distribution in a turbine rotor passage was conducted in a large-scale, ambient temperature, rotating turbine model. Heat transfer was measured for both the full-span suction and pressure surfaces of the airfoil as well as for the hub endwall surface. The objective of this program was to document the effects of flow three-dimensionality on the heat transfer in a rotating blade row (vs a stationary cascade). Of particular interest were the effects of the hub and tip secondary flows, tip leakage and the leading-edge horseshoe vortex system. The effect of surface roughness on the passage heat transfer was also investigated. Midspan results are compared with both smooth-wall and rough-wall finite-difference two-dimensional heat transfer predictions. Contour maps of Stanton number for both the rotor airfoil and endwall surfaces revealed numerous regions of high heat transfer produced by the three-dimensional flows within the rotor passage. Of particular importance are regions of local enhancement (as much as 100 percent over midspan values) produced on the airfoil suction surface by the secondary flows and tip-leakage vortices and on the hub endwall by the leading-edge horseshoe vortex system.

Blair, Michael F.↗
NASA NTRS1972

Exploratory flight investigation of aircraft response to the wing vortex wake generated by jet transport aircraft

The effect of intercepting wing tip vortices generated by large jet transports, including jumbo jets, over separation distances from 1 nautical mile to 15 nautical miles is evaluated on the basis of the response of a vortex probe airplane in the roil mode. The vortex probe test aircraft included a representative general aviation airplane, an executive jet, a fighter, and light and medium weight jet transports. The test conditions and airplane configurations were comparable to those normally used during takeoff, landing, or holding pattern operations. For flight safety the tests were performed at altitudes from 9500 feet to 12,500 feet. In addition to an evaluation of the probe airplane response, a flight test technique is suggested for determining minimum separation distance, using as variable the ratio of vortex-induced roll acceleration to maximum lateral control acceleration and the gross weight of the generating aircraft.

Andrews, W. H.↗
NASA NTRS2005

Control of Flap Vortices

A wind tunnel investigation was carried out on a semi-span wing model to assess the feasibility of controlling vortices emanating from outboard flaps and tip-flaps by actively varying the degree of boundary layer separation. Separation was varied by means of perturbations produced from segmented zero-efflux oscillatory blowing slots, while estimates of span loadings and vortex sheet strengths were obtained by integrating wing surface pressures. These estimates were used as input to inviscid rollup relations as a means of predicting changes to the vortex characteristics resulting from the perturbations. Surveys of flow in the wake of the outboard and tip-flaps were made using a seven-hole probe, from which the vortex characteristics were directly deduced. Varying the degree of separation had a marked effect on vortex location, strength, tangential velocity, axial velocity and size for both outboard and tip-flaps. Qualitative changes in vortex characteristics were well predicted by the inviscid rollup relations, while the failure to account for viscosity was presumed to be the main reason for observed discrepancies. Introducing perturbations near the outboard flap-edges or on the tip-flap exerted significant control over vortices while producing negligible lift excursions.

Greenblatt, David↗
NASA NTRS1990

The acoustic results of a United Techologies scale model helicopter rotor tested at DNW

An initial summary is presented of the acoustic measurements acquired for some of the different configurations of a 1/6 geometrically and aeroelastically scaled UTC model helicopter rotor which was tested in the open-jet anechoic test section of the Duits-Nederlandse Windtunnel in the Netherlands. Of particular interest are high-speed impulsive noise and blade-vortex interaction. An analysis is provided of baseline swept tip rotor acoustic characteristics in the regimes of high-speed forward flight, where high-speed impulsive noise dominates, and low-speed descent, where severe blade vortex interaction noise occurs. Also discussed are more recent studies of data which involve the animation of the acoustic field upstream of the rotor to evaluate the detailed radiation patters caused by BVI and HSI noise sources. The trends of these primary noise sources are examined as the first step in validating the data for release and application.

Liu, Sandy R.↗
NASA NTRS2015

Visualization and Quantification of Rotor Tip Vortices in Helicopter Flows

This paper presents an automated approach for effective extraction, visualization, and quantification of vortex core radii from the Navier-Stokes simulations of a UH-60A rotor in forward flight. We adopt a scaled Q-criterion to determine vortex regions and then perform vortex core profiling in these regions to calculate vortex core radii. This method provides an efficient way of visualizing and quantifying the blade tip vortices. Moreover, the vortices radii are displayed graphically in a plane.

Quantification↗
NASA NTRS1972

Study of induced load and stress, volume 3

An analytical and experimental investigation into the effects of blade tip clearance on inducer performance and of leading edge sweepback on both blade pressure loading and performance was performed. Tip clearance flow was represented with a vortex flow model and measured data from previous inducer tests at three clearances were correlated with model predictions. A leading edge model was added to an existing inducer internal flow analysis, tests with two sweepbacks were conducted, and blade pressure and performance predictions were correlated with measured data.

Coons, L. L.↗
NASA NTRS1977

Measurements of helicopter rotor tip vortices

This paper presents results from a recent wind-tunnel investigation of model helicopter rotor tip vortices. Measurements were made of the vortex positions, core sizes, and velocity distributions. A laser velocimeter was used to make the measurements, and a minicomputer-based data system was used to process the data and to aid in controlling the experiment. The velocimeter, the data system, and the software developed for the minicomputer are briefly described. The rotors investigated were two-bladed, teetering rotors with diameters of 2.1 m. Two sets of blades were used, one set with zero twist and one set with -11 deg of linear twist. The vortex positions were obtained by making flow field traverses while strobing the data system at a fixed azimuth. Aging of a vortex element was also studied by following the convected element while strobing the data system at different azimuths. By this method, the effects on the vortex of a close interaction with a blade and another vortex were studied.

Biggers, J. C.↗