Aircraft noise installation effects

Abstract

Airframe noise is currently of a comparable level to engine noise for an aircraft on approach with high-lift devices and landing gears deployed. The landing gears are a large contributor to the overall airframe noise in this situation. Main landing gears are typically installed beneath a lifting wing. The wing surfaces act as scattering surfaces for the noise generated by these landing gears, and the non-uniform flow around the wing affects both the propagation and strength of the noise. This thesis focuses on investigating the propagation and scattering of installed landing gear noise sources.

Boundary element methods are capable of computing acoustic scattering by large and complex geometries, such as a complete aircraft geometry. However, due to their use of Green’s functions, flow effects can only be approximated. As a result, the refraction of acoustic waves due to a non-uniform flow is not accounted for. A uniform flow formulation based on a Lorentz-type transform is typically employed with boundary element methods. The effect of neglecting refraction on the propagation and scattering of landing gear noise sources is determined in this thesis. Investigations are conducted using computational aeroacoustic methods that solve the linearised Euler equations, which account for the refraction of acoustic waves due to non-uniform flow.

Using computational aeroacoustic methods, the effect of non-uniform flow due to circulation on the acoustic scattering is quantified as the difference in acoustic scattering over uniform and non-uniform base flows. These investigations are conducted using both single frequency and broadband monopole sources, and both single-element and multi-element airfoils. Increasing the angle of attack, increasing the Mach number, and deploying flaps all increase the circulation around the airfoil. The effect of varying these parameters is investigated systematically. It is shown that for a source in the approximate position of a landing gear with flow conditions similar to that of an airliner on approach, the largest difference observed is at single frequencies for an airfoil configuration with a deployed flap. Otherwise, the differences are small, and in some cases so small that they can be considered negligible. It is shown that moving the source to a position above the airfoil and using a higher Mach number gives a larger difference, although this is not representative of a landing gear source.

A new method is proposed to generate a broadband input signal for use with a computational aeroacoustic solver that gives a specified power spectral density at a given radial distance from a monopole source. A signal that is equal in power across a specified range of frequencies is generated using this method. The effect on the frequency content of the scattered noise from a broadband source installed beneath a lifting wing is investigated using this generated signal. It is shown for a single-element airfoil that the major contributor to the obtained power spectral density is the distance of the source from the airfoil. Varying the angle of attack and Mach number has only a small additional effect on the power spectral density. It is then shown that flap and slat deployment has a larger effect on the computed power spectral density due to the additional reflective surfaces.

Existing boundary element method formulations that estimate uniform and nonuniform flow effects are evaluated for their suitability for landing gear noise scattering predictions. It is shown that the uniform flow formulation is more suitable due to a simplifying assumption made in the derivation of the non-uniform flow formulation. An existing realistic landing gear noise model is coupled with a three-dimensional acoustic boundary element method solver. The landing gear noise model applies scaling laws to directional databases for isolated landing gear components in order to estimate the total far-field noise. The implemented coupling methodology is used to compute the sound pressure level on a ground plane beneath a realistic scattering aircraft geometry. The geometrical effect of flap deployment is investigated using sources of constant strength for each configuration. It is shown that the effect of flap deployment is to increase the sound pressure level directly below and in the region immediately surrounding the aircraft. The effect of source strength reduction due to circulation around a lifting wing is then included in the predictions. This results in a large decrease in the predicted sound pressure level on the ground plane with flap deployment.

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