直昇機旋翼葉片所產生之噪音可分為四種主要型態：寬波段噪音、轉動噪音、葉片與渦漩交互作用噪音及高速脈衝噪音。為提升直昇機在懸浮下旋翼葉片之空氣動力性能，利用有限體積法、Roe通量差分分離法及四階Runge-Kutta時間積分法，求解旋轉座標系統下之三維尤拉方程式。模擬NACA0012翼剖面之雙葉片旋翼流場，探討次音速及穿音速之渦流及噪音現象。評估本數值方法結合可調網格之準確性，在次音速方面，利用兩種不同表面網格產生法，在流場區域建立非結構四面體網格與Cardonna實驗值對照。穿音速方面，使用13.71倍旋長之旋翼葉片在三個不同翼尖馬赫數探討噪音聲壓對照於Purcell實驗值；以上驗證均有很好之結果。有效捕捉震波、渦漩和聲波之能力使用可調網格加密技巧，最後在三維流場區域不同截面處之等壓力線圖、等聲壓值圖、渦漩圖及旋翼表面，探討流場渦漩與葉片交互作用及高速震波現象。

　The noise generated by a helicopter rotor blade can be classified into four major types: broadband noise, rotational noise, blade/vortex interaction noise and high-speed impulsive noise. A three-dimensional inviscid flow solver is developed for predicting the aerodynamic performance of hovering helicopter rotor blades using unstructured meshes. The flow solver utilizes a cell-centered finite-volume scheme that is based on the Roe’s flux-difference splitting with Runge-Kutta time integration in a rotating coordinate system. Calculations are performed at two operating conditions of subsonic and transonic tip Mach numbers. A solution-adaptive mesh refinement technique is adopted to improve the resolution of flow features on the blade surface. To accurately simulate the shock, the tip vortex and the related noise, an adaptive mesh refinement technique is adopted. The flowfields for a two-bladed rotor having the NACA0012 airfoil section in hover are studied, and the computed results are compared with those in the related literatures to evaluate the present solution procedure. The near-field Euler uses a solution-adaptive grid scheme to improve the resolution of the acoustic signal. The error indicator is computed from the flowfield solution and determines the regions for mesh coarsening and refinement. Computed results for high-speed impulsive noise compare favorably with experimental data for three different hovering rotor cases. Good agreements are obtained between the numerical result and the experiment. Finally, the contours of pressure and sound pressure level are presented to understand the phenomena of the rotor blade noise.

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