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Investigation of wave phenomena on a blunt airfoil with straight and serrated trailing edges
Abstract
An investigation of pressure waves in compressible subsonic and transonic flow around a generic airfoil is performed in a modified shock tube. New comprehensive results are presented on pressure waves in compressible flow. For the first time, the influence of trailing edge serration will be examined in terms of the reduction in pressure wave amplitude. A generic airfoil is tested in two main configurations, one with blunt trailing edges and the other one with serrated trailing edges in a Mach number range from 0.6 to 0.8 and at chord Reynolds numbers of 1 × 106 < Re
c < 5 ×106. The flow of the blunt trailing edge is characterized by a regular vortex street in the wake creating a regular pattern of upstream-moving pressure waves along the airfoil. The observed pressure waves lead to strong pressure fluctuations within the local flow field. A reduction in the trailing edge thickness leads to a proportional increase in the frequency of the vortex street in the wake as well as the frequency of the waves deduced from constant Strouhal number. By serrating the trailing edge, the formation of vortices in the wake is disturbed. Therefore, also the upstream-moving waves are influenced and reduced in their strength resulting in a steadier flow. An increasing length of the saw tooth enhances the three dimensionality of the structures in the wake and causes a strong decrease in the wave amplitude.
... They found that TE serrations were able to achieve noise reduction of 13 dB in the narrowband frequency due to the reduction of vortex shedding at the TE. The flow around a generic airfoil with blunt and serrated TEs in compressible subsonic and transonic conditions were investigated by Juliane et al. [9] through experimental method. Their results showed that for free stream Mach number up to 0.79, the serrations reduced the pressure wave amplitude significantly and the amount of reduction was proportional to the serration length. ...
... They found that TE serrations were able to achieve noise reduction of 13 dB in the narrowband frequency due to the reduction of vortex shedding at the TE. The flow around a generic airfoil with blunt and serrated TEs in compressible subsonic and transonic conditions were investigated by Juliane et al. [9] through experimental method. Their results showed that for free stream Mach number up to 0.79, the serrations reduced the pressure wave amplitude significantly and the amount of reduction was proportional to the serration length. ...
Two propeller blades were fabricated and tested in an anechoic chamber for thrust force and noise measurements
in order to investigate the effectiveness of serrations on reducing propeller trailing edge noise.
Half flat tip serrations were applied at the trailing edge of the propeller from 0.6R until the tip of the blade.
The serrated propeller had the best performance at 3000 rpm and it obtained the highest insertion loss at
172 Hz which was about 29.6 dBA. By overall, the serrated propeller obtained the highest noise reduction
at 3000 rpm which was about 5.8 dBA. The loss of the thrust force decreased gradually with the increasing
of the rotating speed where the serrated propeller lost about 27.1% of its thrust force at 1500rpm.
This paper presents the results of an experimental investigation exploring the noise reduction potential of sawtooth trailing edge serrations on a flat plate at low-to-moderate Reynolds number (1.6e5 < Rec < 4.2e5). Acoustic and aerodynamic measurements have been taken using a flat plate with both sharp and serrated trailing edges in the anechoic wind tunnel at the University of Adelaide. Trailing edge serrations are found to achieve up to 13 dB of attenuation in the narrowband noise levels without modifying the directivity of the radiated noise. The noise reduction achieved with trailing edge serrations is found to be dependent on their geometrical wavelength and Strouhal number, St = fd/U, where f is frequency, d is boundary layer thickness and U is free-stream velocity. Far-field acoustic data are compared with theoretical noise reduction predictions showing that significant differences exist between measurements and theory. Velocity data measured in the very near trailing edge wake with hot-wire anemometry are related to the far-field noise measurements to give insight into the trailing edge serration noise reduction mechanism. The results suggest that for this particular configuration, the noise reduction capability of trailing edge serrations is related to their influence on the hydrodynamic field at the source location.
Acoustic field measurements were carried out on a 94-m-diam three-bladed wind turbine with one standard blade, one blade with trailing-edge serrations, and one blade with an optimized airfoil shape. A large horizontal microphone array, positioned at a distance of about one rotor diameter from the turbine, was used to locate and quantify the noise sources in the rotor plane and on the individual blades. The acoustic source maps show that for an observer at the array position, the dominant source for the baseline blade is trailing-edge noise from the blade outboard region. Because of convective amplification and directivity, practically all of this noise is produced during the downward movement of the blade, which causes the typical swishing noise during the passage of the blades. Both modified blades show a significant trailing-edge noise reduction at low frequencies, which is more prominent for the serrated blade. However, the modified blades also show tip noise at high frequencies, which is mainly radiated during the upward part of the revolution and is most important at low wind speeds due to high tip loading. Nevertheless, average overall noise reductions of 0.5 and 3.2 dB are obtained for the optimized blade and the serrated blade, respectively.
Exploratory wind-tunnel experiments in high-subsonic and transonic flow on a conventional airfoil with oscillating flap and a supercritical airfoil oscillating in pitch are described. In the analysis of the experimental results, emphasis is placed upon the typical aspects of transonic flow, namely the interaction between the steady and unsteady flow fields, the periodical motion of the shock waves and their contribution to the overall unsteady airloads. Special attention is paid to the behaviour of the supercritical airfoil in its "shock-free" design condition. Moreover, it is discussed to what extent linearization of the unsteady transonic flow problem is allowed if the unsteady field is considered as a small perturbation superimposed upon a given mean steady-flow field. Finally, the current status of unsteady transonic flow theory is reviewed and the present test data are used to evaluate some of the recently developed calculation methods.
Textbook introducing the physical principles and theoretical basis of acoustics, concentrating on concepts and points of view that have proven useful in applications such as noise control, underwater sound, architectural acoustics, audio engineering, nondestructive testing, remote sensing, and medical ultrasonics. Includes problems and answers.
Owls are commonly known for their quiet flight, enabled by three adaptions of their wings and plumage: leading edge serrations,
trailing edge fringes and a soft and elastic downy upper surface of the feathers. In order to gain a better understanding
of the aeroacoustic effects of the third property that is equivalent to an increased permeability of the plumage to air, an
experimental survey on a set of airfoils made of different porous materials was carried out. Several airfoils with the same
shape and size but made of different porous materials characterized by their flow resistivities and one non-porous reference
airfoil were subject to the flow in an aeroacoustic open jet wind tunnel. The flow speed has been varied between approximately
25 and 50m/s. The geometric angle of attack ranged from −16° to 20° in 4°-steps. The results of the aeroacoustic measurements,
made with a 56-microphone array positioned out of flow, and of the measurements of lift and drag are given and discussed.
Die vorliegende Arbeit befasst sich mit umfangreichen experimentellen und numerischen Untersuchungen zur Aerodynamik von Miniature Trailing-Edge Devices (MiniTEDs) an Profilen und Tragflügeln bei transsonischen Anströmgeschwindigkeiten. Für stationäre und instationäre Betrachtungen werden die Einflüsse von MiniTEDs auf die zweidimensionale Profilumströmung und deren aerodynamische Beiwerte für die Gurney Klappe, die Spreizklappe und die divergente Hinterkante detailliert dargestellt. Eine Diskussion der Einflüsse unterschiedlicher Geometrieparameter erfolgt für die Höhe der Gurney-Klappe, den Ausschlagwinkel der Spreizklappe, die MiniTED-Positionierung sowie für Perforationen. Ein variabler MiniTED-Einsatz wird für die widerstandsoptimale Anwendung vorgestellt und der Leistungsfähigkeit einer flexiblen Hinterkante gegenübergestellt. Eine Diskussion der Einflüsse von Anströmparametern erfolgt für Machzahl, Anstellwinkel und Reynoldszahl. Abschließend wird der Einsatz von MiniTEDs an einer realistischen Transportflugzeugkonfiguration dargestellt und hinsichtlich der Beeinflussung von Flügelumströmung und Beiwerten, dem Einfluss der Reynoldszahl sowie einer Validierung der Vorhersage des MiniTED Einflusses bei Flugbedingungen diskutiert und mit Flugversuchsdaten verglichen.
Die vorliegende Arbeit befasst sich mit umfangreichen experimentellen und numerischen Untersuchungen zur Aerodynamik von Miniature Trailing-Edge Devices (MiniTEDs) an Profilen und Tragflügeln bei transsonischen Anströmgeschwindigkeiten. Für stationäre und instationäre Betrachtungen werden die Einflüsse von MiniTEDs auf die zweidimensionale Profilumströmung und deren aerodynamische Beiwerte für die Gurney Klappe, die Spreizklappe und die divergente Hinterkante detailliert dargestellt. Eine Diskussion der Einflüsse unterschiedlicher Geometrieparameter erfolgt für die Höhe der Gurney-Klappe, den Ausschlagwinkel der Spreizklappe, die MiniTED-Positionierung sowie für Perforationen. Ein variabler MiniTED-Einsatz wird für die widerstandsoptimale Anwendung vorgestellt und der Leistungsfähigkeit einer flexiblen Hinterkante gegenübergestellt. Eine Diskussion der Einflüsse von Anströmparametern erfolgt für Machzahl, Anstellwinkel und Reynoldszahl. Abschließend wird der Einsatz von MiniTEDs an einer realistischen Transportflugzeugkonfiguration dargestellt und hinsichtlich der Beeinflussung von Flügelumströmung und Beiwerten, dem Einfluss der Reynoldszahl sowie einer Validierung der Vorhersage des MiniTED-Einflusses bei Flugbedingungen diskutiert und mit Flugversuchsdaten verglichen.
Upstream moving pressure waves are observed in transonic flows over airfoils already for decades. They can be generated actively by flap or airfoil oscillations [6]. But, they are also naturally present in airfoil flows [1, 4]. Upstream moving pressure waves can lead to flow instabilities by forcing shock and stagnation point oscillations. Furthermore, it is expected that they affect the laminar-turbulent transition [5]. Hence, the phenomenon is of great engineering interest.
In vielen technischen Anwendungen treten wegen der geringen ViskositätsWerte Strömungen mit sehr hohen Reynolds-Zahlen auf. Wie in den Beispielen des vorigen Kapitels gezeigt wurde, stellt daher die Grenzlösung Re = ∞ eine gute Näherung dar. Ein schwerwiegender Mangel dieser Grenzlösung ist jedoch, daß sie die Haftbedingung nicht erfüllt, d.h. bei ihr sind die Geschwindigkeiten an der Wand nicht null, sondern endlich. Die Viskosität muß daher berücksichtigt werden, um die Haftbedingung erfüllen zu können. Sie sorgt für den Übergang der Geschwindigkeit vom endlichen Wert der Grenzlösung in Wandnähe zum Wert null direkt an der Wand. Dieser Übergang erfolgt bei großen Reynolds-Zahlen in einer dünnen wandnahen Schicht, die nach L. Prandtl (1904) als Grenzschicht oder auch Reibungsschicht bezeichnet wird. Wie noch gezeigt wird, ist die Dicke der Grenzschicht um so geringer, je größer die Reynolds-Zahl, d.h. je kleiner die Viskosität ist.
Modern civil transport aircraft cruise in the high transonic velocity region near the speed of sound. Mach numbers of up to 0.86 and Reynolds numbers based on the chord length of the wing of up to Rec = 50 × 106 and higher are operational for jet planes like the Airbus A340–600 and the Boeing 747–400. Even bigger aircraft are in the conceptual or design phase with higher passenger capacities e.g. the A380, resulting in larger wing chords to lift the increasing aircraft weights in order to accomodate more passengers or freight per aircraft. Another market trend tends to higher cruising Mach numbers to shorten flight time. So flight Reynolds and Mach numbers will increase in the future.
At the Shock Wave Laboratory (SWL) of the RWTH Aachen University a new wind tunnel was put into operation to perform airfoil testing, duplicating true flight Mach and Reynolds numbers simultaneously. After the calibration phase of the tunnel, extensive airfoil testing has been started. A reproducible test flow for stationary airfoil testing is achieved. Pressure measurements of high temporal resolution and flow visualization are conducted. The tunnel flow in the test section reaches Reynolds numbers of up to Rec = 40 × 106 at Mach numbers between Ma∞ = 0.6 and 0.9.
For the method of a Direct Numerical Simulation (DNS), a mesh study of the transonic flow around the well known NACA 0012 airfoil at a moderate Reynolds number of Rec=5·105 is presented. The three-dimensional Navier-Stokes equations for an unsteady, compressible flow are discretized in a generalized curvilinear coordinate system. The spatial derivatives of first-order are approximated by a fifth-order WENO scheme, the second-order derivatives by a sixth-order central scheme and the time derivatives by a fourth-order Runge-Kutta scheme. The focus of the investigation is on the demonstration of the applicability of DNS for simulating an airfoil flow at a moderate Reynolds number and to study upstream running pressure waves around the airfoil. At this Reynolds number, for a full resolution of all turbulent length scales, theoretically estimated numbers of mesh points are far away from realizable mesh sizes. In the present study seven three-dimensional meshes are compared where each of the two largest meshes consists of one billion mesh points (4096 × 512 × 512 and 8192 × 512 × 256). This mesh size is close to the practical limit of recent simulations since the numerical effort is about 16 · 106 core-hours on a supercomputer for one simulation. The other meshes are gradually coarsened resulting in only four million mesh points for the coarsest mesh. The mesh study is performed by the comparison of aerodynamical and turbulent quantities. On the one hand the main flow features are studied, which are mostly determined by large flow scales. Pressure waves are studied for all meshes, which are generated at the trailing edge, moving upstream. These pressure waves are analyzed in the vicinity of the airfoil. Acoustic phenomena in the far field are not studied. For the present study, a mesh with 67 million mesh points (M3) was sufficient to resolve the main flow features and flow phenomena caused by the pressure waves in the vicinity of the airfoil. On the other hand the turbulent intensities are compared, which are influenced by the smallest turbulent scales. The analysis of the wall units show that even the finest mesh spacings are slightly too large to fulfil the requirements of a fully-resolved DNS. In this context, the energy spectrum of the turbulent kinetic energy is useful to evaluate the quality of the turbulent boundary layer.
Trailing-edge brushes on a supercritical airfoil are investigated in order to analyze their influence on upstream-moving pressure waves evolving from the trailing edge. The upstream-moving pressure waves show a highly-instationary behavior which significantly depends on the local flow velocity on the airfoil. The pressure waves are suspected to have an influence on the flow properties of the airfoil. The current results show their influence on the shock formation. Trailing-edge brushes are installed in order to reduce the amplitude of the upstream-moving pressure waves. By damping the pressure waves a reduction of dynamical loads on the airfoil becomes feasible. Experiments are preformed with a modified shock tube on the supercritical BAC 3-11 airfoil model in a Mach number range from 0.6 up to 0.8 and at a chord Reynolds number of 106. High-speed schlieren photography and pressure measurement are used to observe and quantify the strength of the waves.
Shock tube experiments using differential interferometry have been carried out to investigate compressible subsonic flows around and behind flat plates and cylinders. Pictures taken with an IMACON image converter camera with framing speeds of up to 2.106 frames per second and streak records show that the trailing edge vortex separation induces pressure waves moving upstream. Each vortex separation is related to the formation of a pressure wave. The flow around the bodies is therefore influenced by the shedding frequency and is possibly coupled with the wake. The interaction of the wake with the flow around the bodies is a phenomenon of fundamental importance which has been neglected in gas dynamics up to now. After a symmetrical onset, the process of vortex separation and pressure wave generation develops into an asymmetrical wake with alternating separation. The experiments have shown that the dynamics of vortices close to the trailing edge is practically inviscid and incompressible. In contrast to this, the pressure waves moving upstream must be treated as a problem of purely compressible nature.
This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien-Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack.
Trailing edge (TE) noise measurements for a NACA 63-215 airfoil model are presented, providing benchmark experimental data for a cambered airfoil. The effects of flow Mach number and angle of attack of the airfoil model with different TE bluntnesses are shown. Far-field noise spectra and directivity are obtained using a directional microphone array. Standard and diagonal removal beamforming techniques are evaluated employing tailored weighting functions for quantitatively accounting for the distributed line character of TE noise. Diagonal removal processing is used for the primary database as it successfully removes noise contaminates. Some TE noise predictions are reported to help interpret the data, with respect to flow speed, angle of attack, and TE bluntness on spectral shape and peak levels. Important findings include the validation of a TE noise directivity function for different airfoil angles of attack and the demonstration of the importance of the directivity function’s convective amplification terms.
Wind tunnel measurements on the self-noise of a series of airfoils and flat plates were performed to explore the previously reported noise reducing potential of serrated trailing edges in case of more realistic flows and geometries. For this purpose, different types of airfoils, and flat plates with varying planforms and orientations of the teeth at the trailing edge were used. All serrated airfoils yield reduced trailing-edge noise levels, the reductions ranging from 3 dB up to 8 dB. Spectral shape and dependency on the flow speed and angle-of-attack appeared to be different for every airfoil type. The serrated flat plates were found to give reductions up to 10 dB (1 kHz - 6 kHz). Inclination of the complete flat plate by 10 degrees or a swept trailing edge affected this reduction to a very limited extend only (<2 dB). The same holds for a 10 degrees misalignment of the teeth with respect to the flow direction but in the chord plane. However, misalignment of the teeth by 15 degrees with respect to the chord plane caused an increase of the radiated noise.
Direct numerical simulations of the flow around a NACA-0012 aerofoil are conducted, employing an immersed boundary method to represent flat-plate trailing-edge extensions both with and without serrations. Properties of the turbulent boundary layer convecting over the trailing edge are similar for both cases. For cases with serrations, the trailing-edge noise produced by the flow over the aerofoil is observed to decrease in amplitude, and the frequency interval over which the noise reduction occurs differs depending on the serration length. The directivity and spanwise coherence of the trailing-edge noise appears largely unaffected by the serrations. The hydrodynamic behaviour in the vicinity of the trailing-edge extensions is investigated. The streamwise discontinuity imparted upon the turbulent flow by the straight trailing edge can clearly be observed in statistical quantities, whereas for the serrated case no spanwise homogeneous discontinuities are observed. The trailing-edge serrations appear to break up the larger turbulent structures convecting into the wake, and to promote the development of horseshoe vortices originating at the serrations themselves.
Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by the low-frequency large-amplitude shock motions are highly undesirable from the structural integrity and aircraft maneuverability point of view. For modern supercritical wing design with thick profiles, the shock-induced fluctuations are particularly severe and methods to reduce the shock wave amplitudes to lower values or even to delay the oscillations to higher Mach numbers or incidence angles will result in expanding the buffet boundary of the airfoil. This review begins with a recapitulation of the classical work on shock-induced bubble separation and trailing edge separation of a turbulent boundary layer. The characteristics of the unsteady pressure fluctuations are used to classify the types of shock-boundary layer interaction. The various modes of shock wave motion for different flow conditions and airfoil configurations are described. The buffet boundaries obtained using the standard trailing edge pressure divergence technique and an alternative approach of measuring the divergence of normal fluctuating forces are compared to show the equivalence. The mechanisms of self-sustained shock oscillations are discussed for symmetrical circular-arc airfoils at zero incidence and for supercritical airfoils at high incidence angles with fully separated flows. The properties of disturbances in the wake are examined from linear stability analysis of two-dimensional compressible flows. The advances in high-speed computing make predictions of buffeting flows possible. Navier–Stokes solvers and approximate boundary layer-inviscid flow interaction methods are shown to give good correlation of frequencies and other unsteady flow characteristics with experiments. Finally, passive and active methods of shock oscillation control show promising results in delaying buffet onset to higher Mach numbers or incidence angles, thus enhancing the transonic performance of airfoils.
An analysis is made of the noise produced by low Mach numberturbulent flow over the serrated edge of a flat plate airfoil at zero angle of attack. The serrations are of sawtooth profile of wavelength λ and root‐to‐tip distance 2h. At frequencies ω satisfying ωh/U≫1 (where U is the velocity of the main stream) it is predicted that the intensity of the radiation is reduced relative to that produced by the same flow over an unserrated edge by at least 10×log[1+(4h/λ)2] dB. Predictions are contrasted with analogous results derived [M. S. Howe, J. Fluids Struct. 5, 33–45 (1991)] for smoothly varying serrations of sinusoidal profile, for which it was concluded that attenuations of order 10×log(6h/λ) dB are possible.
This paper provides new results describing compressible fluid flow around a cylinder. The investigation was restricted to subsonic and transonic flow at Reynolds numbers of about 10**5. The experiments showed that a strong coupling exists between the flow over a cylinder and the vortex street formed in the near wake. The phenomenon was investigated using high-speed visualization synchronized with unsteady pressure measurements. Various coupling regimes were classified and instantaneous pressure distributions were obtained at different times during the vortex street period. From these elements it was possible to deduce the unsteady force.
The effects of midspan discrete vortex injection on the performance of a rectangular wing model were studied in a wind tunnel. Based on this preliminary study, discrete vortex injection, while affecting the wing lift and drag, does not degrade its overall performance to any significant degree. This is particularly so in the high-angle-of-attack range where use of vortex injection for wake turbulence alleviation had been proposed. The investigation confirms and expands previously reported, computationally determined results. This study was of a limited scope; additional aspects of discrete vortex injection warrant investigation.
Transonic flow investigations are performed in a modified shock tube with a rectangular test section. The investigated model is a BAC3-11 airfoil with a constant cord length and a sharp trailing edge. Time-resolved shadowgraphs and schlieren pictures show pressure waves initiated near the trailing edge and propagating upstream, where they become apparently weaker near the leading edge. These wave processes are accompanied by wake fluctuations and vortex generation in the boundary layer. The observed waves are also captured by pressure transducers mounted in the airfoil model. The dominant frequencies range between approximately 0.7 and 1.5 kHz. Using statistical analysis of the pressure histories, wave propagation direction and wave speed are determined. For higher flow Mach numbers, a strong wave/shock interaction is also observed in which the shock, depending on the shock strength, is attenuated and degenerated into compression waves. Copyright © 2008 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc.
The present paper deals with the wake of a 2D body equipped with a drag reduction device. The device is a 3D trailing edge consisting of alternate segments of blunt base and spanwise cavity. The aerodynamic mechanisms acting on the near wake are studied in a water tunnel from schlieren observations by thermally marking large scale structures. The results show that the efficiency of the device is directly related to the presence of longitudinal vortices. An optimization of the shapes in subsonic compressible flow had led to a decrease of more than 40% of the total drag of the profile.
The steady aerodynamic effect of "Miniature Trailing-Edge Devices" (MiniTEDs) on a supercritical airfoil was investigated in transonic flow both experimentally and numerically. The investigations were performed for the Gurney flap, the split flap and the divergent trailing edge, and showed a strong influence on the supersonic flow regime, influencing the lift and drag characteristics. At constant lift coefficient a redistribution of the lift generation from the front to the rear part of the airfoil takes place. Total drag reductions can be achieved by the reduction of wave drag. The influences of the different types of MiniTEDs are shown. The investigation of the influence of geometry parameters showed that the variation of the height of the Gurney flap and of the deflection angle of the split flap have similar effects. Perforation of the MiniTEDs showed only a small influence. The investigation of the influence of the free stream parameters indicated that increases in Mach number and Reynolds number increase the MiniTED eect. A drag-optimal application is possible with adaptive MiniTEDs, and the same performance is achieved by an adaptive Gurney flap and an adaptive split flap. A comparison with a flexible trailing edge could identify performance advantages of the MiniTEDs.
Trailing-edge noise data from an experimental study on a flat-plate and on a two-dimensional NACA 0012-like airfoil are presented. Within extensive measurements in the Aeroacoustic Wind Tunnel Braunschweig various flow-permeable (comb-type) edge modifications were tested with respect to their noise reduction capability. Among numerous design parameters a narrow slit width was identified as major requirement for a low-noise trailing-edge design whereas flexibility of the comb material was found to be not imperative for such edge noise reduction means. Because both boundary-layer tripping and trailing-edge thickness show significant influence on trailing-edge noise, a detailed description of these effects is provided to enable an accurate interpretation of corresponding noise test data.
Within a parametric study on brush-type trailing edge extensions the noise reduction potential of several design concepts was determined. The obtained data base represents the first phase of an ongoing project with the long-term objective to develop scaling laws for a future application of such devices as add-on solutions for today's aircraft components. The experiments comprised both acoustic and aerodynamic measurements on a zero-lift generic plate model (Re = 2.1 to 7.9 x 10 Million) in DLR's open jet Aeroacoustic Wind Tunnel Braunschweig, AWB. Noise data were taken by means of a directional microphone system.
A comprehensive experimental investigation of trailing edge noise is reported for the case of a two-dimensional airfoil embedded in a uniform low Mach number flow. The Reynolds number is high and the boundary layer is fully turbulent. Parameters include angle of attack, flow velocity, and trailing edge bluntness. By using a coherent output power method, the trailing edge noise spectra and directivity (including forward speed effects) are quantitatively determined. Statistics of the pressure field beneath the turbulent boundary layer are defined in detail. The scattered pressure field (primarily incompressive) very near the trailing edge is measured and successfully modeled by extending existing theory. This helps establish the edge condition (of pertinence to the Kutta condition) which in turn determines the solution for the sound field in this study. By using a statistical model of the turbulent boundary layer pressure field, trailing edge noise is well predicted.
A review is made of the diffraction theory of the trailing edge noise generated by a flat-plate airfoil of zero-thickness and non-compact chord, according to which the sound is attributed to the scattering of a “frozen” pattern of turbulence wall pressure swept over the edge in the mean flow. Extension is made to determine the sound produced by very low Mach number flow over the edge of an airfoil of finite thickness. In applications it is desirable to represent the noise in terms of a surface integral over the airfoil involving a Green's function and a metric of the edge flow that can be calculated locally using the equations of motion of an incompressible fluid. It is argued that the appropriate metric for a rigid airfoil is the incompressible “upwash” velocity (determined by the Biot–Savart induction formula applied to the boundary layer vorticity outside the viscous sublayer), and not the surface pressure. Formulae for calculating the noise are given when the airfoil thickness is acoustically compact, and for both three- and two-dimensional edge flows.The theory is illustrated by a detailed discussion of a two-dimensional vortex flow over an airfoil with a rounded trailing edge. The problem is simple enough to be treated analytically, yet is also suitable for validating computational edge noise schemes.
A discussion is given of the production of sound by low Mach number turbulent flow over the trailing edge of a serrated airfoil. The airfoil is modeled by a flat plate set at zero angle of attack to the mean flow, and attention is given to both limiting cases in which the chord of the airfoil is either large or. small relative to the characteristic acoustic wavelength; a formula is proposed for interpolating predictions at intermediate frequencies. General arguments are advanced which imply that, for serrations of spanwise wavelength λ and amplitude h, and at radian frequencies ω satisfying ωh/U ≫ 1 (U being the velocity of the main stream), the frequency spectrum of trailing edge noise is reduced relative to that for an unserrated edge by 10 × log10{Ch/λ} (dB) when λ/h ≲ 4, where the constant C ≅ 10. This conclusion is confirmed, and predictions are extended to include larger values of λ/h, by an approximate analytical treatment of the case involving an edge with sinusoidal serrations, and by use of an empirical model of the turbulent flow. At high frequencies the predicted attenuation is about 1 dB when λ/h ≃ 10; at λ/h = 1 the attenuation is typically 7 or 8 dB. It is argued that, in principle, optimal attentuations should be obtained by use of serrations of sawtooth profile with edges inclined at less than 45° to the direction of the mean flow.
The noise emission generated by the passage of a turbulent airstream over the trailing edge of a semiinfinite plate was measured over a large range of airstream velocity and plate geometry. The experiment was designed to validate trailing-edge noise theories. The results show that the peak of a radiation pattern moves from an upstream to a downstream direction as the velocity increases. The measured radiation pattern of the noise was in agreement with that predicted by a recent fundamental theory for leading- and trailing-edge noise. Although large changes in the character of the turbulent flow near the trailing edge effect the level and spectra of trailing-edge noise, the shape of the pattern is still accurately predicted by this theory.
Determination of the vortex shedding frequency of cascades with different trailing edge thicknesses
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