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Brooks, T. F.

Publications and source records attributed to Brooks, T. F..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2007

Aero-acoustics of Drag Generating Swirling Exhaust Flows

Aircraft on approach in high-drag and high-lift configuration create unsteady flow structures which inherently generate noise. For devices such as flaps, spoilers and the undercarriage there is a strong correlation between overall noise and drag such that, in the quest for quieter aircraft, one challenge is to generate drag at low noise levels. This paper presents a rigorous aero-acoustic assessment of a novel drag concept. The idea is that a swirling exhaust flow can yield a steady, and thus relatively quiet, streamwise vortex which is supported by a radial pressure gradient responsible for pressure drag. Flows with swirl are naturally limited by instabilities such as vortex breakdown. The paper presents a first aero-acoustic assessment of ram pressure driven swirling exhaust flows and their associated instabilities. The technical approach combines an in-depth aerodynamic analysis, plausibility arguments to qualitatively describe the nature of acoustic sources, and detailed, quantitative acoustic measurements using a medium aperture directional microphone array in combination with a previously established Deconvolution Approach for Mapping of Acoustic Sources (DAMAS). A model scale engine nacelle with stationary swirl vanes was designed and tested in the NASA Langley Quiet Flow Facility at a full-scale approach Mach number of 0.17. The analysis shows that the acoustic signature is comprised of quadrupole-type turbulent mixing noise of the swirling core flow and scattering noise from vane boundary layers and turbulent eddies of the burst vortex structure near sharp edges. The exposed edges are the nacelle and pylon trailing edge and the centerbody supporting the vanes. For the highest stable swirl angle setting a nacelle area based drag coefficient of 0.8 was achieved with a full-scale Overall Sound Pressure Level (OASPL) of about 40dBA at the ICAO approach certification point.

Shah, P. N.↗
NASA NTRS2002

An Aeroacoustic Study of a Leading Edge Slat Configuration

Aeroacoustic evaluations of high-lift devices have been carried out in the Quiet Flow Facility of the NASA Langley Research Center. The present paper describes detailed flow and acoustic measurements that have been made in order to better understand the noise generated from airflow over a wing leading edge slat configuration, and to possibly predict and reduce this noise source. The acoustic database is obtained by a moveable Small Aperture Directional Array of microphones designed to electronically steer to different portions of models under study. The slat is shown to be a uniform distributed noise source. The data was processed such that spectra and directivity were determined with respect to a one-foot span of slat. The spectra are normalized in various fashions to demonstrate slat noise character. In order to equate portions of the spectra to different slat noise components, trailing edge noise predictions using measured slat boundary layer parameters as inputs are compared to the measured slat noise spectra.

Mendoza, J. M.↗
NASA NTRS1994

Recent studies of rotorcraft blade-vortex interaction noise

Recent results are presented from several research efforts aimed at the understanding of rotorcraft blade-vortex interaction (BVI) in terms of the noise generation, directivity, and control. The results are based on work performed by NASA Langley Research Center researchers, both alone and in collaboration with other research organizations. Based on analysis of a simplified physical model, the critical parameters controlling BVI noise generation have been identified. The detailed mapping of the acoustic radiation field of a model rotor in a wind tunnel has revealed the extreme sensitivity of directivity to rotor advance ratio and disk attitude. The control and reduction of BVI noise through the use of higher harmonic pitch control is discussed.

BLADE VORTEX INTERACTION↗
NASA NTRS1990

Rotorcraft blade/vortex interaction noise - Its generation, radiation, and control

Recent results are presented from several research efforts aimed at the understanding of rotorcraft blade-vortex interaction noise generation, directivity, and control. The results are based on work performed by researches at the NASA Langley Research Center, both alone and in collaboration with other research organizations. Based on analysis of a simplified physical model, the critical parameters controlling the noise generation are identified. Detailed mapping of the acoustic radiation field reveals the extreme sensitivity of directivity to rotor advance ratio and disk attitude. A means of controlling blade-vortex interaction noise by higher harmonic pitch control is discussed.

Preisser, J. S.↗
NASA NTRS1987

A directional array approach for the measurement of rotor noise source distributions with controlled spatial resolution

A special array system has been designed to examine noise source distributions over a helicopter rotor model. The particular measurement environment is for a rotor operating in the open jet of an anechoic wind tunnel. An out-of-flow directional microphone element array is used with a directivity pattern whose major directional lobe projects on the rotor disk. If significant contributions from extraneous tunnel noise sources in the direction of the side lobes are excluded, the dominant output from the array would be that noise emitted from the projected area on the rotor disk. The design incorporates an array element signal blending features which serves to control the spatial resolution of the size of the directional lobes. (Without blending, the resolution and side lobe size are very strong functions of frequency, which severely limits the array's usefulness).

Brooks, T. F.↗
NASA NTRS1986

Reduction of Vane Noise in Wind-Tunnel Nozzles

Vane-induced noise reduced by adhering thin layer of porous material such as foam to downwind surface of vanes, particularly near sharp trailing edges, to prevent occurrence of edge tones. Other modifications made to aerodynamically streamline vane structure were rounding leading edges and smoothly fairing trailing edges. Boundary layer trip applied to leading edge to prevent laminar tone shedding from trailing edge.

Martin, R. M.↗
NASA NTRS1985

Reduction of background noise induced by wind tunnel jet exit vanes

The NASA-Langley 4 x 7 m wind tunnel develops low frequency flow pulsations at certain velocity ranges during open throat mode operation, affecting the aerodynamics of the flow and degrading the resulting model test data. Triangular vanes attached to the trailing edge of flat steel rails, mounted 10 cm from the inside of the jet exit walls, have been used to reduce this effect; attention is presently given to methods used to reduce the inherent noise generation of the vanes while retaining their pulsation reduction features.

Martin, R. M.↗
NASA NTRS1985

Scaling of airfoil self-noise using measured flow parameters

Data from an airfoil broadband self-noise study are reported. Attention here is restricted to two-dimensional models at zero angle of attack to the flow. The models include seven NACA 0012 airfoil sections and five flat plate sections with chordlengths ranging from 2.54 to 60.96 cm. Testing parameters include flow velocity to 71.3 m/s and boundary-layer turbulence through natural transition and by tripping. Detailed aerodynamic measurements are conducted in the near-wake of the sharp trailing edges. The noise spectra of the self-noise sources are determined by the use of a cross-spectral technique. The acoustic data are normalized using the measured aerodynamic parameters in order to evaluate a commonly used scaling law. An examination of the Reynolds number dependence of the normalized overall levels has revealed a useful scaling result. This result appears to quantify the transition between turbulent boundary-layer trailing-edge noise and laminar boundary-layer vortex shedding noise.

Brooks, T. F.↗
NASA NTRS1984

Airfoil trailing edge flow measurements and comparison with theory, incorporating open wind tunnel corrections

Trailing edge data for boundary layer-near wake thickness parameters are given for airfoils and flat plates. Reynolds number effects are examined as a function of model size, velocity and boundary layer tripping. These data expand that presented previously by the authors particularly for airfoil non-zero angles of attack. Comparisons are made here with boundary layer calculations using potential flow modeling and a well documented two-dimensional finite-difference method for laminar and turbulent boundary layers. Open wind tunnel corrections to angle of attack and camber are developed and are incorporated in the potential flow modeling to assure correct comparisons for non-zero angles of attack. It was found that although the open tunnel flow turbulence affected boundary layer transition for the higher velocities the theory successfully 'brackets' the data. Comparisons demonstrate the degree of accuracy one might expect for the prediction of boundary layer thickness parameters when given only geometry and nominal flow conditions as input to boundary layer codes.

Brooks, T. F.↗
NASA NTRS1984

Airfoil tip vortex formation noise

Spectral data are presented for the noise produced due to the turbulent three-dimensional vortex flow existing near the rounded tip of lifting airfoils. The results are obtained by the comparison of sets of two- and three-dimensional test data for different airfoil model sizes, angles of attack, and tunnel flow velocities. Microphone cross-correlation and cross-spectral methods were used to determine the radiated noise. Corrections were made for tunnel shear layer and source directivity effects. Interpretation of the results are aided by a three-dimensional flow analysis developed for this study which determines open tunnel and finite aspect ratio corrections heretofore neglected in tip vortex studies. Hot wire measurements were made in the tip vortex formation region for the specification of governing flow parameters. The spectral data is normalized in a format considered most useful for subsequent quantitative prediction of this noise mechanism for practical systems such as helicopter rotors. Comparison is made to the analysis of George and Chou. A recommended prediction method is given.

Brooks, T. F.↗
NASA NTRS1983

Airfoil self noise - Effect of scale

Key data from a comprehensive airfoil broadband self-noise study are reported. Attention here is restricted to two-dimensional sharp trailing-edge models. The models include seven NACA 0012 airfoil sections and five 'flat plate' sections with chordlengths ranging from 2.54 to 60.96 cm. Testing parameters include flow velocity, angle of attack to the flow, and boundary layer turbulence through natural transition and through tripping. Detailed aerodynamic measurements, of pertinence to the scaling problem of airfoil self-noise, were conducted in the near-wake of the trailing edges. Presented are mean and rms turbulent velocity profiles as well as boundary layer thicknesses and integral thickness parameters for a large range of conditions. The noise spectra of the self-noise sources were determined by the use of a cross-spectral technique. The spectra were normalized using the measured aerodynamic parameters in order to evaluate the most common scaling law now in use. An examination of the Reynolds number dependence of the overall self-noise levels has revealed a new and useful scaling result. This result appears to quantify the transition between turbulent-boundary-layer trailing edge noise and laminar-boundary-layer vortex shedding noise.

Brooks, T. F.↗
NASA NTRS1983

Progress in rotor broadband noise research

The current status of research on various rotor broadband noise sources is reviewed. Theoretical modeling techniques which appear to have the most promise in the development of successful noise prediction capabilities are emphasized. The types of noise considered include: turbulence ingestion noise, blade self-noise due to turbulence passing the trailing edge, blade self-noise due to separated flow, and blade self-noise due to vortex shedding. Comparisons with experimental results are given, and present theoretical and experimental limitations are delineated to help identify areas of needed research.

Brooks, T. F.↗
NASA NTRS1981

Trailing edge noise prediction from measured surface pressures

Trailing edge (TE) noise is investigated for the case of a two-dimensional airfoil embedded in a uniform low Mach number flow, and the usefulness of several TE noise theories is examined by applying them to the measured data. The TE noise spectra and directivity are quantitatively determined for the case of a high Reynolds number and a fully turbulent boundary layer. Parameters include angle of attack, flow velocity and TE bluntness. Evanescent wave theories by Chase (1975) and Chandiramani (1974) are compared to the present results and show good agreement. Agreement of the near field pressure scatter phenomenon analysis with measurements implies that the basic assumptions used in the analysis are correct, i.e., the turbulent boundary layer (TBL) flow passes the trailing edge into the wake region. No hydrodynamic wake shedding activity is confirmed for the two-sided TBL flow, and a method incorporating the principles of the coherent output power method is used to determine the sound field. The near field edge scatter model is found to not only establish optimum sizing of edge treatment for noise control, but also to separate and identify the scattered field from the incident hydrodynamic field.

Brooks, T. F.↗
NASA NTRS1981

Trailing edge noise prediction using Amiet's method

Amiet's (1976, 1978) solution to the problem of airfoil trailing edge noise prediction is discussed in light of the results of evanescent wave theory's application to the measured surface pressure behavior near the trailing edge of an airfoil with a turbulent boundary layer. The method employed by Amiet has the advantage of incorporating the effect of finite chord in its solution. The assumed form of the pressure distribution is examined as well as the constant turbulent boundary layer convection assumption, which is found to be unnecessarily restrictive.

Brooks, T. F.↗
NASA NTRS1979

An experimental investigation of trailing-edge noise

Airfoil trailing-edge noise up to a Reynolds number based on chord of 2.96 x 10 to the 6th power was studied. Comparisons are made with current theory, particularly with regard to the nature of the pressure field in the vicinity of the trailing-edge and its influence on the radiated noise.

Brooks, T. F.↗
NASA NTRS1979

Investigation of trailing-edge noise

A comprehensive experimental investigation of airfoil trailing-edge noise up to a Reynolds number based on chord of 2.96 x 10 to the 6th power is described. Comparisons are made with current theory, particularly with regard to the pressure field in the vicinity of the trailing-edge and its influence on the radiated noise.

Brooks, T. F.↗