Experiment of the far-field acoustic characteristics of an airfoil with smooth rime ice model on the leading edge
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摘要:
结冰厚度与气动噪声增量的关系为探索一种新型的结冰厚度探测方法提供了新思路。在英国南安普敦大学的低噪声航空声学风洞中开展了前缘带光滑霜冰模型的NACA0012翼型远场噪声特性实验，翼型的远场噪声信号采用环形麦克风阵列测量，远场声压信号利用快速傅里叶变换处理，得到了最大结冰厚度、来流速度和来流攻角对NACA0012翼型远场噪声声压级的影响。结果表明：光滑霜冰模型改变了翼型前缘的局部流场，流动分离导致了远场噪声特性的较大变化。在实验条件下，结冰翼型与基准翼型间远场声压级的最大增量超过9.5 dB，出现在频率8×103～2×104 Hz范围。最大结冰厚度、来流攻角、来流速度与结冰翼型的远场声压级呈正相关性，建立了一种飞行参数、总声压级增量等多变量输入的最大结冰厚度神经网络预测模型。
Abstract:The relationship between ice thickness and aeroacoustic increment provides a new idea for exploring a new method of ice thickness detection. The far-field acoustic characteristics of a NACA0012 airfoil with smooth rime ice model on the leading edge were studied experimentally. The low-noise aeroacoustic wind tunnel of the University of Southampton was used as the experimental platform. The far-field noise signals were measured by a arc microphone array. The far-field sound pressure was processed by fast Fourier transform to obtain the effects of the maximum ice thickness, airflow speed and angle of attack on the sound pressure level of NACA0012 airfoil. The results showed that, the smooth rime ice model changed the local flow field around the leading edge of the airfoil, resulting in a relatively large variation of far-field acoustic characteristics by flow separation. The maximum increment of the far-field sound pressure level between iced airfoil and baseline exceeded 9.5 dB within the frequency between 8×103−2×104 Hz. There was a positive correlation between the maximum ice thickness, airflow speed, angle of attack and the far-field sound pressure level of iced airfoil. A neural network prediction model of maximum icing thickness with multivariable inputs such as flight parameters and overall sound pressure level increment was established.
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Key words:
- ice accretion /
- airfoil /
- leading edge /
- far-field /
- acoustics /
- wind tunnel experiment /
- sound pressure level
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图 1 航空声学风洞和环形麦克风阵列的实验设置
Figure 1. Sketch of aeroacoustic wind tunnel and arc microphone array setup
图 3 前缘带光滑霜冰模型的NACA0012翼型示意图
Figure 3. Sketch of NACA0012 airfoil with smooth rime ice on leading edge
图 4 安装于开口实验段的实验模型
Figure 4. Experimental model installed in the open experimental section
图 6 不同最大结冰厚度翼型远场声压级（U=40，60，80 m/s，α=0°）
Figure 6. Far-field sound pressure level of airfoil with different maximum ice thicknesses （U=40，60，80 m/s，α=0°）
图 7 不同最大结冰厚度翼型总声压级分布（U=40，60，80 m/s，α=0°）
Figure 7. Over sound pressure level distribution of airfoil with different maximum ice thicknesses （U=40，60，80 m/s，α=0°）
图 8 不同来流攻角下翼型远场声压级（h=4.54 mm，U=40，60，80 m/s）
Figure 8. Far-field sound pressure level of airfoil with different angles of attack （h=4.54 mm，U=40，60，80 m/s）
图 9 不同来流攻角下翼型总声压级分布（h=4.54 mm，U=40，60，80 m/s）
Figure 9. Over sound pressure level distribution of airfoil with different angles of attack （h=4.54 mm，U=40，60，80 m/s）
图 10 不同来流速度下翼型远场声压级（h=0，2.29，4.54 mm，α=0°）
Figure 10. Far-field sound pressure level of airfoil with different airflow speeds （h=0，2.29，4.54 mm，α=0°）
图 11 不同来流速度下翼型总声压级分布（h=0，2.29，4.54 mm，α=0°）
Figure 11. Over sound pressure level distribution of airfoil with different airflow speeds （h=0，2.29，4.54 mm，α=0°）
图 13 不同最大结冰厚度翼型附近的涡量云图（U=60 m/s，α=10°）
Figure 13. Vorticity contour around airfoil with different maximum ice thicknesses （U=60 m/s，α=10°）
图 14 结冰翼型产生气动噪声机理的示意图
Figure 14. Sketch of generation mechanism for aeroacoustic from airfoil on the leading edge
图 15 结冰翼型绕流流场和流线图（h=4.54 mm，α=0°，5°，10°）
Figure 15. Flowfield and streamline around airfoil with ice model （h=4.54 mm，α=0°，5°，10°）
表 1 3种预测模型预测值与实际值的相对误差
Table 1. Relative errors between practical value and predicted value of three prediction models
总声压级增量/dB 来流速度/（m/s） 来流攻角/（°） 神经网络误差/% 线性回归误差/% 非线性回归误差/% 3.5 40 0 −0.00026 −4.25787 −71.02521 2.3 80 0 −0.00024 −18.96472 −24.52846 2.0 60 5 0.00020 10.06066 −14.53293 3.2 80 5 −0.00003 −20.69472 −23.20168 2.6 60 10 −0.02143 −3.65963 −13.95765 
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