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Buzz-saw noise is a main component of the noise measured in the forward-arc of transonic fans during the take-off and climbing of aircraft. Full resolving of this noise component is computationally expensive because it is very sensitive to random blade variations, such as manufacturing tolerance, deformations or defects. To avoid this, artificially-synthesized irregular buzz-saw noise source is widely used for the prediction and further study. However, the irregularities of the synthesized source are hardly linked to the blade geometry and the so-induced flow structures. In this study, it is found that blade stagger, axial and azimuth variations maintain a quasi-linear relation with the so-induced static pressure perturbations. Based on this quasi-linear relation, the disturbances induced by blade variations are represented by a set of acoustic modal perturbations generated by several pairs of "basic" rotors. By superposing such modal perturbations, the buzz-saw noise source and the static pressure field generated by a fan with multiple arbitrary blade variations can be achieved instantly. This method is programmed to develop a code package named the Mode-based Fast Reconstruction (MFR). It is validated from simple to complex application scenarios based NASA rotor 67. An accuracy level of 0.9 dB is achieved for the multiple pure tones generated by rotors with multiple blade variations that satisfy the quasi-linear assumptions. This enables high-accuracy fast reconstruction of buzz-saw noise source and pressure field in more real transonic fans. Additionally, the impact of blade variations on buzz-saw noise is much stronger than the influence on rotor aerodynamic performances. So, the effects of small blade variations should be considered during the design and optimization of transonic fan rotors, even though the aerodynamic performances may not benefit much from this. [ABSTRACT FROM AUTHOR]

Most of the published works related to noise generation in high-lift devices only consider clean idealized geometries, whereas real aircraft wings contain important parts called the slat tracks that connect the slat and the main wing. To investigate the effects of slat tracks on the ambient flow and noise, simulations are performed for wings based on the 30P30N airfoil. Two configurations are considered, i.e., the 30P30N wing with and without a slat track. Improved detached-eddy simulation is carried out to compute the unsteady near-field flow. The far-field noise propagation is calculated using the Ffowcs Williams-Hawkings equation. The results show that the installation of the slat track has observable effects on the aerodynamic and aeroacoustic performances of the high-lift devices. The maximum lift coefficient of the high-lift device and the stall angle of attack is reduced. Low-frequency broadband noise radiated from the main wing is significantly increased and the directivity of the overall sound pressure level is changed due to the slat track. The far-field noise from the configuration with the slat track displays noticeable sensitivity to the angle of attack. The slat track has an amplification effect on the high-pressure fluctuation on the main wing, especially at a small angle of attack. This leads to nearly 10 dB in the far-field noise based on the discussed wing section. To reveal the relative noise mechanism, vortical structures near the slat track, large-scale spanwise flows above the main wing, and the vortex dynamics are investigated. A spiral vortex pair is captured as the most important flow structure. It is generated in the spanwise direction by the interaction of slat wakes with the boundary layer on the upper surface of the main wing. The pressure fluctuations on the main wing are highly correlated with this flow structure, which cannot be ignored for the study of noise in a real aircraft. [ABSTRACT FROM AUTHOR]

The noise of slat in the high-lift device of large aircraft is one of the main noise sources of large aircraft during take-off and landing. This paper investigates a noise reduction measure which consists in droop nose for high-lift device of three-dimensional sweep wing instead of leading-edge slat. The unsteady flow field and the far-field noise are predicted by a hybrid SA-IDDES method coupled with the integration of the Ffowcs Williams–Hawkings equation (FW-H). Based on the optimal droop nose configuration from 2D airfoil, the results of the 3D swept wing model with the droop nose on the leading edge of the inner wing are given. Numerical results show that the droop nose configuration can reduce the overall sound pressure level (OASPL) of the intermediate frequency noise, and the peak noise can be reduced by 15 dB, and the sound pressure level of the broadband noise at low and high frequencies is also partially suppressed. The maximum lift-to-drag ratio of the droop nose configuration increased by 2.7 %. In addition, the noise reduction mechanism of leading-edge droop is analyzed by combining the vortex sound theory and the generation and evolution process of the vortex structure. The leading-edge droop nose configuration fundamentally not only eliminates the vortex structure of the slat cove of the leading-edge slat configuration, but also avoids the interaction between the slat slot flow and the upper surface of the main wing. These directly lead to a thinner boundary layer vortex near the top of the main wing, and the wall pressure pulsation is smaller than that of the leading-edge slat high-lift device, thereby reducing noise. The reduction of the BEF distribution on the upper surface boundary shows again that the droop nose inhibits the vortex on the upper surface of the main wing. The change of the Lamb vector divergence reveals the noise reduction mechanism of the drooping leading edge configuration directly from the sound source. [ABSTRACT FROM AUTHOR]

A novel wall treatment method, termed as the Slat Cove Wavy Wall (SCWW), is proposed to improve both the aero-acoustic and the aerodynamic performances of high-lift devices. By this treatment, a wavy pattern is imposed on the wall in the slat cove. In the present work, slats with three different SCWWs are designed and applied to a straight wing composed by 30P30N airfoil. Numerical simulations and analyses are conducted to unveil the underlying mechanisms of noise reduction. The unsteady flow field and the far-field noise are predicted by a hybrid SA-IDDES method followed by the integration of the Ffowcs Williams and Hawkings equation. The modified high-lift device is proved to be able to improve the aero-acoustic performance by reducing the low-frequency components of the radiated noise. Meanwhile, the aerodynamic performance is maintained or improved under proper design parameters. The boundary layer separation in the slat cove is accelerated by the SCWW, which transforms low-frequency periodic vortex shedding to high-frequency turbulent flow. This results in the sound energy shift from low-frequency noise to high-frequency broadband noise. The SCWW method is proved to be effective in redistribution of boundary vorticity flux on the slat lower surface. To reveal the mechanism of the SCWW's drag and noise reduction, the vortex dynamic processes near the slat lower surface in terms of the Lamb vector divergence are further investigated. It is found that the Lamb vector divergence is directly related to both the noise source and the drag. The interaction area of positive and negative Lamb vector divergence can be balanced and decreased by SCWW, resulting in the drag and noise reduction. As expressed by the vortex dynamic processes, this feature implies the SCWW method is of great potential in noise attenuation. [ABSTRACT FROM AUTHOR]