1,721,006 research outputs found
A weak-scattering model for tone haystacking caused by sound propagation through an axisymmetric turbulent shear layer
No full textIn aeroacoustics, spectral broadening refers to the scattering of tonal sound fields by turbulent shear layers, whereby the interaction of the sound with turbulent flow results in power lost from the tone and distributed into a broadband field around the tone frequency. Fan and turbine tone spectral broadening is known colloquially as “haystacking”. Spectral broadening causes substantial difficulties in determining the relative importance of various sources of turbofan engine noise, as it scatters sound generated by tonal sources across a wide frequency range. This obscures the importance of some key sources of tone noise by apparently reducing their prominence, the resultant broadband noise being often difficult to attribute. Recently a new analytical solution has been derived to predict weak spectral broadening of a tone radiated through a circular jet. In this paper, validation of the weak scattering model is presented. A key aspect of the modelling is the choice of the two-point turbulent velocity cross-correlation function which is used to provide a statistical description of the turbulence in the shear layer. Results obtained using three different cross-correlation functions are compared. These include cross-correlation functions which have been developed using the theory for homogeneous isotropic turbulence or using the theory for homogeneous axisymmetric turbulence. In particular, a new cross-correlation function for an axisymmetric turbulent shear layer formed by a circular jet, based on the theory for homogeneous axisymmetric turbulence, has been developed. Validation results of weak-scattering calculated using this correlation function show better agreement with measurements when compared to the results calculated using correlation functions based on the theory for three-dimensional homogeneous isotropic turbulence
Spectral broadening of tonal sound propagating through an axisymmetric turbulent shear layer
No full textIn aeroacoustics, spectral broadening refers to the scattering of tonal sound fields by turbulent shear layers, whereby the interaction of the sound with turbulent flow results in power lost from the tone and distributed into a broadband field around the tone frequency. Fan and turbine tone spectral broadening is known colloquially as “haystacking”. Recently a new analytical model has been derived to predict weak spectral broadening of a tone radiated through a circular jet. A key part of the modeling is the choice of the two-point turbulent velocity cross-correlation function, which is used to provide a statistical description of the turbulence in the shear layer. A new cross-correlation function for an axisymmetric turbulent shear layer formed by a circular jet, based on the theory for homogeneous axisymmetric turbulence, has been developed. Validation results of weak-scattering calculated using this correlation function show better agreement with measurements when compared with the results calculated using a correlation function based on the theory for homogeneous isotropic turbulence
Propeller tone scattering
No full textBecause of recent uncertainty in the price of fuel and also because of concerns over carbon emissions from aircraft, aeronautical engine manufacturers are exploring various alternative propulsors to the turbofan engine. One such propulsor is the advanced open rotor - which consists of two coaxial, counter-rotating propellers, and which promises a significant fuel burn reduction relative to current generation turbofan engines. However, in order for the open rotor to become a viable propulsor it must meet stringent noise emission targets. Therefore, having noise prediction tools which allow the quick and accurate assessment of the noise produced by a large number of advanced open rotor designs are a necessary part of the design process. Frequency domain formulae are often used for this purpose. These formulae contain a Fourier integral in the coordinate coaxial with the propeller axis, which in the near-field must be evaluated numerically at a high computational cost. For far-field noise predictions, and if it is assumed that the propeller operates in an environment with no scattering surfaces, then the Fourier integral can be evaluated in a relatively straight-forward manner, and at low computational cost, using the method of stationary phase. Consequently, most of the published literature regarding propeller and advanced open rotor noise ignores scattering from surrounding aircraft structures. In this paper methods for including the effect of scattering from structures such as the fuselage and the open rotor hub are presented. Some results showing the effect of scattering on the radiated sound field are also presented.<br/
Prediction of fan tone radiation scattered by a cylindrical fuselage
No full textA theoretical prediction method of the scattering of fan tone radiation from a turbofan inlet duct by the airframe fuselage is presented. The fan tone noise is modelled by an acoustic disc source that represents the sound field at the inlet duct termination. Adjacent to the source is a cylindrical fuselage that scatters the fan tone radiation. The prediction method is valid for upstream sound radiation. The acoustic pressure on the cylindrical fuselage is affected by refraction of the sound as it propagates through the fuselage boundary layer. This effect known as boundary layer shielding is more prominent forward of the turbofan, since the fan tone noise radiated from the inlet duct is propagating upstream. An asymptotic approach is used to model sound propagation through a boundary layer which is modelled by a thin linear shear velocity profile. Consequently the scattered pressure field can be computed very quickly, thus providing a fast and efficient prediction method. Although a realistic fuselage turbulent boundary layer does not resemble a linear shear layer, it is shown that the effect of acoustic shielding by a turbulent boundary layer can be modelled by taking a liner shear profile with a shape factor that matches the shape factor for a realistic turbulent profile
An analytical model of sound refraction by the fuselage boundary layer for fan tone radiation from a turbofan aero-engine
No full textThe work presented in this paper is on the development and validation of an analytical-based model to predict scattering of fan tone noise from a turbofan engine by the airframe fuselage and refraction due to the presence of a boundary layer on the fuselage. The aim is to avoid numerical solutions to calculate sound propagation through the fuselage boundary layer which have been prevalent in previous work on this topic. The work presented here is a continuationof the work by the authors in which the 1/7th power-law boundary-layer velocity profile was replaced with a linear profile which enabled an analytical solution of the Pridmore-Brown equation. The work presented here offers an even simpler alternative solution by replacing the power-law profile with a step function. This approach once again leads to a far-field solution in terms of a Fourier series, and a near-field solution expressed in terms of a Fourier series and an inverse Fourier transform. The two approaches utilising a linear or step-function velocity profile are compared with each other, and against existing numerical results. The results show that for sufficiently thin boundary-layers both approaches can approximate a more realistic power-law boundary-layer profile, while the accuracy of the step-function solution appears to not deteriorate with thicker layers. A parametric study based on multi-modal simulations is performed with realistic operating and flight conditions. Predictions for boundary-layer shielding and the far-field polar directivity are generated with both linear and step-functio
Theoretical methods for the prediction of near-field and far-field sound radiation of fan tones scattered by a cylindrical fuselage
No full textThe aim of the work presented in this paper is the development of theoretical methods to predict scattering of fan tone noise from a turbofan engine by the airframe fuselage. The analysis begins with an overview of previous research on fan tone noise scattering by an adjacent cylindrical fuselage. In all similar previous work the propagation of sound through the fuselage boundary layer has been calculated using numerical methods. The effect of the boundary layer can be very significant on the upstream radiated sound from a turbofan’s intake. An asymptotic approach is presented to model sound propagation within the boundary layer.An entirely analytic formulation is derived for a thin linear velocity profile. This approach leads to a far-field solution expressed in terms of a Fourier series, and a near-field solution expressed in terms of a Fourier series and a Fourier inverse transform. The new formulation is validated by comparison with simpler analytic solutions, and against existing numerical solutions. Furthermore, the results using a linear velocity profile are shown to be comparable with numerical results calculated with a realistic shear velocity profile that closely matches a turbulent boundary layer. Preliminary results from the new theoretical method are presented that illustrate the refraction effect by the fuselage boundary layer
A theoretical model of fuselage pressure levels due to fan tones radiated from the intake of an installed turbofan aero-engine
No full textAn existing theoretical model to predict the pressure levels on an aircraft's fuselage is improved by incorporating a more physically realistic method to predict fan tone radiation from the intake of an installed turbofan aero-engine. Such a model can be used as part of a method to assess cabin noise. Fan tone radiation from a turbofan intake is modelled using the exact solution for the radiated pressure from a spinning mode exiting a semi-infinite cylindrical duct immersed in a uniform flow. This approach for a spinning duct mode incorporates scattering/diffraction by the intake lip, enabling predictions of the radiated pressure valid in both the forward and aft directions. The aircraft's fuselage is represented by an infinitely long, rigid cylinder. There is uniform flow aligned with the cylinder, except close to the cylinder's surface where there is a constant-thickness boundary layer. In addition to single mode calculations it is shown how the model may be used to rapidly calculate a multi-mode incoherent radiation from the engine intake. Illustrative results are presented which demonstrate the relative importance of boundary-layer shielding both upstream and downstream of the source, as well as examples of the fuselage pressure levels due to a multi-mode tonal source at high Helmholtz number
Fuselage boundary-layer refraction of fan tones radiated from an installed turbofan aero-engine
No full textA distributed source model to predict fan tone noise levels of an installed turbofan aero-engine is extended to include the refraction effects caused by the fuselage boundary layer. The model is a simple representation of an installed turbofan, where fan tones are represented in terms of spinning modes radiated from a semi-infinite circular duct, and the aircraft’s fuselage is represented by an infinitely long, rigid cylinder. The distributed source is a disc, formed by integrating infinitesimal volume sources located on the intake duct termination. The cylinder is located adjacent to the disc. There is uniform axial flow, aligned with the axis of the cylinder, everywhere except close to the cylinder where there is a constant thickness boundary layer. The aim is to predict the near- field acoustic pressure, and in particular to predict the pressure on the cylindrical fuselage which is relevant to assess cabin noise. Thus no far-field approximations are included in the modelling. The effect of the boundary layer is quantified by calculating the area-averaged mean square pressure over the cylinder’s surface with and without the boundary layer included in the prediction model. The sound propagation through the boundary layer is calculated by solving the Pridmore-Brown equation. Results from the theoretical method show that the boundary layer has a significant effect on the predicted sound pressure levels on the cylindrical fuselage, owing to sound radiation of fan tones from an installed turbofan aero-engine
Prediction of Swirl Effects on Fan-OGV Interaction Tones
No full textNoise levels predicted for the fan-OGV interaction tones generated by modern high and ultra high bypass ratio aircraft engines are significantly changed when swirl in the interstage region is included in the modelling. An analytical prediction method is used to predict interstage interaction tonal noise levels for an annular duct with uniform flow or swirling flow. It is predicted that swirl effects alter the range of modes that are cut-on, and their corresponding sound power levels in the upstream direction. Results confirm that the inclusion of swirl effects in the modelling of fan-OGV interaction tones is important to improve the prediction of the sound power levels of the upstream propagating modes
Generation of discrete frequency tones by the flow around an aerofoil
Full text linkTonal noise, the self-induced discrete frequency noise generated by aerofoils, is investigated. It is heard from an aerofoil placed in streams at low Mach number flows when inclined at a small angle to the stream. The tones are heard as a piercing whistle, commonly up to 30 dB above the background noise level. The work is motivated by the occurrence of tonal noise from rotors, fans and recently wind-turbines. Previous authors have attributed tonal noise to a feedback loop consisting of a coupling between laminar boundary-layer instability waves and sound waves propagating in the free stream. The frequency has been predicted by use of various methods based on this model.In this thesis a review of wind-tunnel results obtained by Dr. E.C. Nash at the University of Bristol is presented. Boundary-layer measurements show the presence of tonal noise is closely related to the existence of a region of separated flow close to the trailing edge of the aerofoil. Highly amplified boundary-layer instability waves were observed close to the trailing edge of the aerofoil at the frequency of the tone. A comprehensive analysis of the linear stability of the boundary-layer flow over the aerofoil is presented. The growth of boundary-layer instability waves over theaerofoil is calculated. The growth rates of the waves were obtained by solving the Orr–Sommerfeld problem at several stations on the aerofoil. The Falkner–Skan boundary layers were found to be a suitable form of velocity profiles to incorporate the adverse pressure gradients experienced by the flow over an aerofoil. The amplification of the instability waves is shown to be controlled almost entirely by the region of separated flow close to the trailing edge. The calculated frequency of the linear modes with maximum amplification over the aerofoil is found to be close to the observed frequency of the acoustic tone.A weakly nonlinear stability analysis was also performed and this appears to be a suitable description of the boundary-layer instability waves. The results indicate that the frequency of the tones may commonly be predicted to within 10% by using weakly nonlinear stability theory.The generation of sound by diffraction of the boundary-layer instability waves at the trailing edge of the aerofoil is also discussed as well as the proposed feedback models. A modified feedback model is proposed, being based on the experimental and theoretical results
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