To investigate the vortex dynamics and loading noise radiation characteristics during dynamic stall of helicopter rotor airfoils, a numerical study is carried out using the unsteady Reynolds-averaged Navier–Stokes method and the Ffowcs Williams-Hawkings equation. First, the aerodynamic and aeroacoustic characteristics of NACA0012 airfoil in dynamic stall were calculated. The numerical method is validated through comparisons with experimental data. Then, the aeroacoustic characteristic of an airfoil in the typical deep stall condition is analyzed, with a detailed discussion on the relationship between flow phenomena and sound pressure (SP). Based on the vortex evolution over airfoil surface, a model for loading noise driven by vortex evolution during dynamic stall is presented. Finally, the impact of reduced frequency (k) on the loading noise characteristics of the airfoil is analyzed, and conclusions are obtained. The dynamic stall loading noise is primarily caused by the shedding and movement of the dynamic stall vortex (DSV) and trailing edge vortex, with DSV shedding inducing rapid and intense SP fluctuations, resulting in high-intensity noise of up to approximately 30 Pa. An increase in k enhances the rate of change in unsteady surface pressure coefficient (Cp), which accelerates pressure variation and leads to a significant rise in overall sound pressure level.

Based on the Ffowcs Williams–Hawkings equations and the computational fluid dynamics method, the rotor’s aeroacoustic characteristics, considering the influence of the downwash flowfield on the sound propagation process, are calculated and analyzed. First, a set of analysis methods for the aeroacoustic characteristics is developed, and a convection-based propagation time model is developed, where acoustic group velocity along source–observer lines quantifies flowfield effects. Then, the rotor’s aerodynamic and aeroacoustic characteristics are calculated, and the employed numerical analysis method is validated through the comparisons with experimental data. Finally, the aeroacoustic characteristics of the rotor in hover are analyzed, and the sound pressure positive peak point with/without the influence of the flowfield of downwash on the propagation time is discussed in detail. In addition, parameters, such as the rotor’s collective pitch and the azimuthal angle of the sound source, are quantified, and some conclusions are obtained. For those observers below the rotor rotation plane, the downwash flowfield will influence the sound propagation time, resulting in the increase in the sound pressure and the advance of the arrival time.

For high-speed compound helicopters, such as the S-97 Raider, the reflection and diffraction effects of vertical/horizontal tails on pusher propeller noise are inevitable. To investigate the noise distortion effect of the rear-mounted pusher propeller, this study first relies on the Chinese Laboratory of Rotorcraft Navier-Stokes (CLORNS) solver, adopting the high-resolution Perturbed polynomial reconstructed Targeted Essentially Non-Oscillatory scheme (TENO-P) combined with the Delayed Detached Eddy Simulation based on the Spalart–Allmaras (SA-DDES) turbulence model to resolve the multi-scale rotor flowfield. Additionally, a continuous and conserved acoustic source extraction method is proposed to eliminate non-physical waves at the one-way Computational Fluid Dynamics and Computational AeroAcoustics (CFD–CAA) coupling interface, addressing the temporal inconsistency between flowfield evolution and acoustic propagation. Finally, numerical investigations are conducted on the instantaneous acoustic wave propagation and acoustic directivity of the pusher propeller under the influence of vertical/horizontal tails. The results show that significant acoustic distortion occurs when pusher propeller-generated noise interacts with vertical/horizontal tails. This interaction not only produces reflected and diffracted acoustic waves but also leads to wavefront discontinuities, the formation of short acoustic waves, and changes in acoustic directivity. The maximum variation in the sound pressure level reaches 10 dB at local azimuths. The distortion effect of tails on pusher propeller noise is closely correlated with the number of propeller blades. The interaction process between the propeller and tails becomes more complex with the increase in blade count, resulting in the generation of shorter acoustic waves. For the six-blade rotor, the originally continuous acoustic wave branch can be split into up to four short waves. This study confirms that the proposed Hybrid Computational AeroAcoustics (HCAA) method holds significant application prospects in the aeroacoustic research of compound helicopters.