“…For the case with 0.88 blade tip Mach number, the microphone at 1.136 times the radius of the blade is within the acoustic source region (as shown in Figure 14), so the acoustic result obtained by the CFD method is presented (as shown in Figure 19 (c)). The CFD result shows excellent agreement with the experimental data, 52 and this indicates that the CFD result is quite reliable as the input for the CAA solver. The results of the HCAA method for the other three in-plane microphones are presented in Figure 19(d), (e), and (f), and the results agree better to the experimental data than the results of the FW-H_pds (Cy) method, while its peak negative pressure for the microphone at 3.09 times the radius of the blade is slightly underpredicted than the FW-H_pds (Ad) method due to the dispersion and the dissipation of the numerical method.Figure 19.In-plane sound pressure-time history for high-speed impulsive noise.…”
“…The near-field acoustic characteristic of the UH-1H model rotor 51,52 is calculated based on the HCAA method. This rotor is a one-seventh scale of the UH-1H main rotor with straight and untwisted blades with a NACA0012 airfoil section.…”
Section: Validation Of the Noise Prediction Methods Under Uniform Flo...mentioning
confidence: 99%
“…The high-speed impulsive noise is caused by the compressibility effect of rotor flowfield when the blade tip Mach number is large enough. Some excellent methods have been developed to predict the high-speed impulsive noise, such as the purely CFD method based on Euler equations 51 or the Ffows Williams and Hawkings equations with adaptive integration surface 52 (abbreviated as FW-H_pds (Ad)). The FW-H_pds (Ad) method is developed based on the fact that the placement of the integration surface affects the results of the FW-H_pds equation.…”
“…The FW-H_pds (Ad) method is developed based on the fact that the placement of the integration surface affects the results of the FW-H_pds equation. Two criteria are presented by Chen et al 52 for determining the integration surface. The two criteria are related to density perturbation and pressure gradient respectively.…”
“…The CFD grid size within the CFD-CAA coupling surface and the CAA grid size in Zone 1 (Zone 1 in Figure 3) should be small enough to reduce the dispersion and dissipation of the acoustic waves. The high-speed impulsive noise data of the UH-1H model rotor is used to verify the current HCAA method, 51,52 and the cases with blade tip Mach numbers of 0.8, 0.85, and 0.88 are analyzed below. As illustrated in Figure 13, the case with 0.88 blade tip Mach number is the most difficult one to be simulated, and the side length of the tetrahedron for this case should be less than 0.09 Ch.…”
To study the influence of non-uniform flowfield on the propagation characteristics of helicopter rotor noise, a Hybrid Computational Aeroacoustics (HCAA) method is developed. The acoustic source region is simulated by Computational Fluid Dynamics (CFD) technique with the Unsteady Reynolds Averaged Navier-Stokes equations (URANS) as the governing equations. Acoustic near-field is simulated by Computational Aeroacoustics (CAA) technique with the Linearized Euler Equations (LEE) as the governing equations, and the numerical discretization of the LEE is accomplished by Runge-Kutta Discontinuous Galerkin (RKDG) method. A novel acoustic source extraction method based on pressure and pressure gradient is proposed to accomplish the one-way CFD-CAA weak coupling. The HCAA method is validated through comparisons with noise experimental data of the UH-1H model rotor and the BO-105 model rotor. Based on the proposed HCAA method, the convection and refraction effects of rotor noise under different collective pitch angles are analyzed. The results show that the distortion effect of the rotor noise is most affected by the non-uniformly distributed downwash velocity field, resulting in an increment of acoustic energy below the rotor plane. The effect of non-uniformly distributed downwash velocity on noise propagation increases with the increase of the collective pitch angle. For the UH-1H model rotor, the maximum change of the sound pressure level is 0.8 dB (about 10% change of the effective sound pressure).
“…For the case with 0.88 blade tip Mach number, the microphone at 1.136 times the radius of the blade is within the acoustic source region (as shown in Figure 14), so the acoustic result obtained by the CFD method is presented (as shown in Figure 19 (c)). The CFD result shows excellent agreement with the experimental data, 52 and this indicates that the CFD result is quite reliable as the input for the CAA solver. The results of the HCAA method for the other three in-plane microphones are presented in Figure 19(d), (e), and (f), and the results agree better to the experimental data than the results of the FW-H_pds (Cy) method, while its peak negative pressure for the microphone at 3.09 times the radius of the blade is slightly underpredicted than the FW-H_pds (Ad) method due to the dispersion and the dissipation of the numerical method.Figure 19.In-plane sound pressure-time history for high-speed impulsive noise.…”
“…The near-field acoustic characteristic of the UH-1H model rotor 51,52 is calculated based on the HCAA method. This rotor is a one-seventh scale of the UH-1H main rotor with straight and untwisted blades with a NACA0012 airfoil section.…”
Section: Validation Of the Noise Prediction Methods Under Uniform Flo...mentioning
confidence: 99%
“…The high-speed impulsive noise is caused by the compressibility effect of rotor flowfield when the blade tip Mach number is large enough. Some excellent methods have been developed to predict the high-speed impulsive noise, such as the purely CFD method based on Euler equations 51 or the Ffows Williams and Hawkings equations with adaptive integration surface 52 (abbreviated as FW-H_pds (Ad)). The FW-H_pds (Ad) method is developed based on the fact that the placement of the integration surface affects the results of the FW-H_pds equation.…”
“…The FW-H_pds (Ad) method is developed based on the fact that the placement of the integration surface affects the results of the FW-H_pds equation. Two criteria are presented by Chen et al 52 for determining the integration surface. The two criteria are related to density perturbation and pressure gradient respectively.…”
“…The CFD grid size within the CFD-CAA coupling surface and the CAA grid size in Zone 1 (Zone 1 in Figure 3) should be small enough to reduce the dispersion and dissipation of the acoustic waves. The high-speed impulsive noise data of the UH-1H model rotor is used to verify the current HCAA method, 51,52 and the cases with blade tip Mach numbers of 0.8, 0.85, and 0.88 are analyzed below. As illustrated in Figure 13, the case with 0.88 blade tip Mach number is the most difficult one to be simulated, and the side length of the tetrahedron for this case should be less than 0.09 Ch.…”
To study the influence of non-uniform flowfield on the propagation characteristics of helicopter rotor noise, a Hybrid Computational Aeroacoustics (HCAA) method is developed. The acoustic source region is simulated by Computational Fluid Dynamics (CFD) technique with the Unsteady Reynolds Averaged Navier-Stokes equations (URANS) as the governing equations. Acoustic near-field is simulated by Computational Aeroacoustics (CAA) technique with the Linearized Euler Equations (LEE) as the governing equations, and the numerical discretization of the LEE is accomplished by Runge-Kutta Discontinuous Galerkin (RKDG) method. A novel acoustic source extraction method based on pressure and pressure gradient is proposed to accomplish the one-way CFD-CAA weak coupling. The HCAA method is validated through comparisons with noise experimental data of the UH-1H model rotor and the BO-105 model rotor. Based on the proposed HCAA method, the convection and refraction effects of rotor noise under different collective pitch angles are analyzed. The results show that the distortion effect of the rotor noise is most affected by the non-uniformly distributed downwash velocity field, resulting in an increment of acoustic energy below the rotor plane. The effect of non-uniformly distributed downwash velocity on noise propagation increases with the increase of the collective pitch angle. For the UH-1H model rotor, the maximum change of the sound pressure level is 0.8 dB (about 10% change of the effective sound pressure).
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