4, in both internal (channel) and external (airfoil) flows. A review of the boundary-layer parameters characterizing the pressure gradient effects is provided, and the more relevant ones are introduced as new variables in the model. The method is then compared to the zero pressure gradient model it is derived from. The influence of the pressure gradient on the wall-pressure spectrum is discussed.Finally,themethodisappliedtoprovideinputdataofaradiatedtrailing-edge noisemodelbymeansofan aeroacoustic analogy, namely Amiet’s theory of turbulent boundary layer past a trailing edge. The results are compared to experimental data obtained in an open-jet anechoic wind tunnel.

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Wall-Pressure Spectral Model Including the Adverse Pressure Gradient Effects

This paper summarizes the results of an experimental study on flow-permeable trailing-edge noise reduction means. Basic design rules are presented with respect to a future employment at current high-lift devices of passenger aircraft. The main focus is directed at the identification of the major design parameters of comb-type or slit edge-modifications. The achievable noise reduction capability was quantified by directional microphone measurements on a flat plate and on a two-dimensional NACA0012-like airfoil in the open-jet Aeroacoustic Wind Tunnel Braunschweig. It was found that flexibility of the comb material is beneficial, but not essential to achieve a noise reduction. Apart from a minimum device length the slit width was identified as the decisive design parameter. An almost zero-spacing of the comb fibers revealed the best results, leading to the assumption that the obtained noise reduction is mainly due to a viscous damping of turbulent flow pressure amplitudes in the comb area.

Design Criteria for Low-Noise Trailing-Edges

This paper is concerned with the application of porous treatments as a means of flow and aerodynamic noise reduction. An extensive experimental investigation is undertaken to study the effects of flow interaction with porous media, in particular in the context of the manipulation of flow over blunt trailing edges and attenuation of vortex shedding. Comprehensive boundary layer and wake measurements have been carried out for a long flat plate with solid and porous blunt trailing edges. Unsteady velocity and surface pressure measurements have also been performed to gain an in-depth understanding of the changes to the energy–frequency content and coherence of the boundary layer and wake structures as a result of the flow interaction with a porous treatment. Results have shown that permeable treatments can effectively delay the vortex shedding and stabilize the flow over the blunt edge via mechanisms involving flow penetration into the porous medium and discharge into the near-wake region. It has also been shown that the porous treatment can effectively destroy the spanwise coherence of the boundary layer structures and suppress the velocity and pressure coherence, particularly at the vortex shedding frequency. The flow–porous scrubbing and its effects on the near-wall and large coherent structures have also been studied. The emergence of a quasi-periodic recirculating flow field inside highly permeable surface treatments has also been investigated. Finally, the paper has identified several important mechanisms concerning the application of porous treatments for aerodynamic and aeroacoustic purposes, which can help more effective and tailored designs for specific applications.

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Trailing-edge flow and noise control using porous treatments

The present investigation of aerofoil self-noise generation and propagation concerns the effects of mean flow and quadrupole sources on the broadband noise that arises from the interaction of turbulent boundary layers with the aerofoil trailing edge and the tonal noise that arises from vortex shedding generated by laminar boundary layers and trailing-edge bluntness. Compressible large-eddy simulations (LES) are conducted for a NACA0012 aerofoil with rounded trailing edge for four flow configurations with different angles of incidence, boundary layer tripping configurations and free-stream Mach numbers. The Reynolds number based on the aerofoil chord is fixed at . The acoustic predictions are performed by the Ffowcs Williams & Hawkings (FWH) acoustic analogy formulation and incorporate convective effects. Surface and volume integrations of dipole and quadrupole source terms appearing in the FWH equation are performed using a three-dimensional wideband multi-level adaptive fast multipole method (FMM) in order to accelerate the calculations of aeroacoustic integrals. In order to validate the numerical solutions, flow simulation and acoustic prediction results are compared to experimental data available in the literature and good agreement is observed in terms of both aerodynamic and aeroacoustic results. For low-Mach-number flows, quadrupole sources can be neglected in the FWH equation and mean flow effects appear only for high frequencies. However, for higher speeds, convection effects are relevant for all frequencies and quadrupole sources have a more pronounced effect for medium and high frequencies. The convective effects are most readily observed in the upstream direction.

Convective effects and the role of quadrupole sources for aerofoil aeroacoustics

Experimental characterization of the turbulent boundary layer over a porous trailing edge for noise abatement

Wall pressure sources near an airfoil trailing edge under turbulent boundary layers

A stochastic estimation technique has been applied to simultaneously acquired data of velocity and surface pressure as a tool to identify the sources of wall-pressure fluctuations. The measurements have been done on a NACA0012 airfoil at a Reynolds number of Re
 c
  = 2 × 105, based on the chord of the airfoil, where a separated laminar boundary layer was present. By performing simultaneous measurements of the surface pressure fluctuations and of the velocity field in the boundary layer and wake of the airfoil, the wall-pressure sources near the trailing edge (TE) have been studied. The mechanisms and flow structures associated with the generation of the surface pressure have been investigated. The “quasi-instantaneous” velocity field resulting from the application of the technique has led to a picture of the evolution in time of the convecting surface pressure generating flow structures and revealed information about the sources of the wall-pressure fluctuations, their nature and variability. These sources are closely related to those of the radiated noise from the TE of an airfoil and to the vibration issues encountered in ship hulls for example. The NACA0012 airfoil had a 30 cm chord and aspect ratio of 1.

Ana Garcia-Sagrado

Papers

Wall pressure sources near an airfoil trailing edge under turbulent boundary layers

Stochastic estimation of flow near the trailing edge of a NACA0012 airfoil

Experimental Investigation Into Trailing Edge Noise Sources

Wall-Pressure Sources Near an Airfoil Trailing Edge Under Separated Laminar Boundary Layers