INVESTIGATIONS ON SLAT NOISE REDUCTION TECH- NOLOGIES BASED ON PIEZOELECTRIC MATERIAL, PART II: CONTROL SYSTEM DESIGN AND WIND TUNNEL TEST Song Xiao, Yu Jinhai, Breard Cyrille and Sun Yifeng Shanghai Aircraft Design and Research Institute (SADRI), COMAC, Shanghai, China email: sunyifeng@comac.cc Qiu Jinhao and Ji Hongli College of Aerospace Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing, China Aeroacoustic noise from the civil aircraft is not only related to its airworthiness, but also related to the demand of comfort. The noise from slat is the major contribution to the overall airframe noise. Analysis and control of slat noise is of great importance for improving noise environment. A thorough understanding of the flow physics and underlying noise mechanisms is indispensable for prediction, control, and mitigation of slat noise. Two main activities were investigated. Firstly, experiments were performed to investigate the flow field and noise characteristics of a two-element airfoil system. Visualization tests with oil flow, pressure and velocity measurements were carried out in order to investigate the flow characteristic of slat cove. The broadband component and multiple tonal peaks of slat noise were identified by sound intensity technology. Secondly, based on the understanding of the unsteady physics mechanism of the flow in slat cove, a method using active vibration excitation for slat noise control was proposed and validated. The control system was installed at the cusp of slat using piezoelectric (PZT) material in order to disturb the vortex shedding. This paper identifies the slat noise by experiment and reveals active vibration excitation on slat noise reduction. The experimental results show the multiple tonal peaks of slat noise are induced by the feedback mechanism of cavity, and application of the vibration excitation can break the feedback mechanism of slat cove. The proposed control method has an impact on multiple tonal peaks of slat noise and a 4dB reduction in the fourthorder tonal peak was achieved. 1. Introduction With the development of the high-bypass ratio engines in the last decades, aircraft engine noise has come down to a level comparable to airframe noise [1]. Both wind tunnel tests [2][3] and flyover measurements [4][5] have shown that leading-edge slat represents a major source of airframe noise during approach and landing. In recent years, more and more scholars have devoted their efforts to the research of slat noise reduction. Several active and passive control methods have been made on the attenuation of slat noise, but there is an ongoing debate on the application of these technologies. In this paper, the flow physics and underlying noise mechanisms of a three-element airfoil system are summarized in detail. Based on this, a method of active vibration control with piezoelectric fans is introduced and its effectiveness is verified by simulative results both in aspects of aerodynamics and aeroacoustics. 1
2. Noise mechanics Slat noise mainly consists of low-frequency broadband noise and high-frequency tonal noise. Major slat noise sources are depicted in figure 1. The high-frequency component is related to the vortex shedding from the finite thickness edge, making contribution to the high-frequency waves. It should be noted that this phenomena is not likely observed at a real full scale model due to the smaller relative edge thickness comparative to a two-dimensional scale model experiments. The low-frequency component is the most prominent component in slat noise spectra in a real aircraft. Although its mechanisms haven t been well understood, a more acceptable view [6] is as follows. The free shear layer sheds off the slat cusp before it rolls up and forms several large-scale coherent structures. An evident train of vorticity structures along the slat shear layer ends on the slat cove surface, where strong vortex-body and vortex-vortex interactions occur. Flow reattachment of free shear layer vortex is identified for the potential broadband source mechanism. Comparative to vortex shedding from bluntness, it makes contribution to the low-frequency waves. As is generally agreed, slat noise is broadband in nature. Therefore, efforts in slat noise reduction must focus on the low-frequency broadband component. vortex shedding secondary separation vortex impingement edge scattering 3. Experimental design Figure 1: Slat noise source mechanisms. In this subsection, we introduce the test model and wind tunnel, then, discuss the new noise control method. 3.1 Configurations The test model is designed by Shanghai Aircraft Design and Research Institute design for slat noise research, as shown in figure 2. The airfoil takes cavity flow characteristics of 30P30N [7] for reference. Compared to three element airfoil, the two element airfoil is easier to be installed in wind tunnel, and has no trailing edge flap noise. Figure 2: Two dimensional airfoil. 3.2 Wind Tunnel The wind tunnel is a low speed wind tunnel of Nanjing University of Aeronautics and Astronautics, as shown in Figure 3. In this wind tunnel, a special wind tunnel controller (FSCG05/P05 CONVO) is designed, which is mainly used to provide a stable flow field for the aerodynamic test of various kinds of models. Open test section size: 0.8m (L) * 0.4m (W) * 0.6m (H). Motor power is less than 11KW. Adjusting the motor power, we can obtain the wind speed we expected in the 2 ICSV23, Athens (Greece), 10-14 July 2016
test section. The wind uses soft connection between the wall of the wind tunnel and the motor so as to reduce the influence of the mechanical vibration of the motor. Figure 3: Wind tunnel. Aerodynamic noise research is generally carried out in an open wind tunnel, because the closed test section of the wall has a strong acoustic reflection and the sound insulation effect. The test model installation method is shown in Figure 4. The slat chord length is 0.088m and the slat span length is 0.6m. In the wind tunnel outlet, a collapsible organic glass plate is fixed to form a flow guiding plate. Figure 4: The installation mode of model. 3.3 Control methods The method of active vibration control is briefly described in figure 5. In figure 6, a slotted wall with two rows of piezoelectric fans are designed to replace aircraft skin, the location and dimension of the slot are also marked. Each piezoelectric fan is fabricated by bonding a piezoelectric patch to a shim material. The piezoelectric fan is simplified as a cantilever beam system.the idea behind the above method is to disturb the formation of shear layer by a local unsteady force, expecting a reduction of vortex impingement intensity on the slat cove surface. Figure 5: The control method. ICSV23, Athens (Greece), 10-14 July 2016 3
4. Results and Discussion Figure 6: The design of piezoelectric actuator. In this subsection, we discuss aerodynamic and aeroacoustic characteristics of baseline configuration, then, focus on the control effect of low noise configuration. 4.1 Aerodynamic aspects The flow field is measure by hot wire anemometer and oil flow. The hot wire is a precision measuring instrument, which can convert wind speed to electric signal, and then receive the wind speed value through the data collector. Velocity distribution in the cavity is obtained by measuring the velocity of each point in the slat cove by a hot wire velocity meter. The wind speed is 20m/s. The measurement method is shown in figure 7. Figure 7: Measurement using hot wire anemometer. The velocity image of slat cavity is the flow velocity of the intermediate XY plane. It is unable to measure the velocity distribution in the vicinity of the reattached area and the seam path. Figure 8 shows the experimental measurement of the velocity of the slat cavity. The interface between the low velocity zone and the external flow field is the free shear layer, the measurement by the hot wire identifies the approximate location of the shear layer. 4 ICSV23, Athens (Greece), 10-14 July 2016
0.00-0.02 y[m] -0.04-0.06-0.08 0.02 0.04 0.06 0.08 0.10 x[m] Figure 8: Velocity distribution of the cavity. The surface oil flow technique is applied to oil flow on the wall, it can visual the wall flow characteristics through the movement of the tracer particles in the coating. The wind speed is 20m/s. As shown in Figure 9, when the air impact the wall of the cavity wall, it leaves a white line, which is the reattachment point position. The distance between trailing edge and the reattachment point position is about 21mm Figure 9: Experimental results of oil flow. 4.2 Aeroacoustic aspects Sound intensity measurement has the advantages of strong anti-interference ability, and is not affected by the background noise. It is an effective method for noise source identification and sound field analysis. The test wind tunnel is a common low speed wind tunnel, not an acoustic wind tunnel. Environmental noise and background noise of wind tunnel are inevitable during the experiment, so the sound intensity method is chosen to measure the radiated noise. As shown in Figure 10, the sound intensity probe is located outside the wind tunnel test section, which is opposite to the point of attachment in the middle plane. The wind speed is 15m/s, 20m/s and 25m/s. The sampling frequency is 8192Hz, the sampling number is 4096, the resolution of frequency domain analysis is 2Hz. The linear average of each sample, the average number of times is 100 times. Figure 11: Sound intensity measurement. ICSV23, Athens (Greece), 10-14 July 2016 5
Figure 11 compares the point of the A with and without test model (wind tunnel background). It can be seen that the background noise decays from low frequency to high frequency, and the background noise is larger in the 20Hz~200Hz range. In this frequency band, the noise is submerged in the background noise, which can not be recognized. In the 200Hz~2000Hz range, the noise is much higher than the background noise of the wind tunnel. The noise obviously reduces with the increase of the frequency above 2000Hz. The results show that the 200Hz~2000Hz band is a low frequency broadband noise band, but it is difficult to be identified because of the interference of the measurement environment. PSD[dB/Hz] 110 100 90 80 70 60 50 40 30 20 10 100 1000 Frequency[Hz] without model with model 300Hz~2000Hz Figure 12: The sound intensity of point A with/without model. 4.3 Control effect The original model is defined as baseline configuration, the model is defined as low noise configuration after the installation of the actuator. Figure 12 compares the noise intensity results of the baseline configuration and the low noise configuration. As can be seen in the range of 200Hz~1500Hz, far field noise of low noise configuration declines about 10dB, which further illustrates shielding part recirculation zone can reduce the far field radiation noise of slat. 120 100 baseline configuration low noise configuration PSD[dB/Hz] 80 60 40 20 0 100 1000 Frequency[Hz] Figure 13: Sound intensity of baseline configuration and low noise configuration at point A. By applying a voltage of bimorph, we study the influence of actuator vibration on the slat noise. The reduction of point A is mainly concentrated in the fourth order (880Hz), and the other order of the peak is almost unchanged. It can be seen that the active control method can control the far field noise, but only reduces the fourth order, whose energy is relatively weak. 6 ICSV23, Athens (Greece), 10-14 July 2016
90 80 0Hz 600Hz PSD[dB/Hz] 70 60 50 40 500 1000 1500 Frequency[Hz] Figure 14: Control effect. 5. Conclusions An active vibration control method applied to slat noise reduction is studied in this paper. The introduction of the vibrational surface changes the configuration. The deformation of the vibrational surface is very small, but it changes the shear layer stemming from the cusp. The results, achieved by experiment, show a 4 db reduction of farfield noise. Current works have verified the effectiveness of the active vibration control. Further work will concern on the concept of optimum parameters. REFERENCES 1 Dobrzynski W. Almost 40 years of airframe noise research: What did we achieve? Journal of aircraft, 47(2): 353-367, (2010). 2 Choudhari M M, Lockard D P, Macaraeg M G, et al. NASA/TM-2002-211432, Aeroacoustic experiments in the Langley low-turbulence pressure tunnel, (2002). 3 Takeda, K., Zhang, X., & Nelson, P. A. Unsteady aerodynamics and aeroacoustics of a high-lift device configuration, 40 th AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, U.S.A., (2002). 4 Olson, S., Thomas, F., & Nelson, R. Mechanisms of slat noise production in a 2D multi-element airfoil configuration, 7 th AIAA/CEAS Aeroacoustics Conference and Exhibit. Maastricht, Netherlands, (2001) 5 Mendoza, J. M., Brooks, T. F., & Humphreys, W. Aeroacoustic measurements of a wing/slat model, 8 th AIAA/CEAS Aeroacoustics Conference & Exhibit. Breckenridge, CO U.S.A., 17-19 June, (2002) 6 Choudhari, M. M., & Khorrami, M. R. Effect of three-dimensional shear-layer structures on slat cove unsteadiness. AIAA journal, 45(9), 2174-2186, (2007). 7 Klausmeyer S M, Lin J C. Comparative results from a CFD challenge over a 2D three-element high-lift airfoil. National Aeronautics and Space Administration, Langley Research Center, (1997). ICSV23, Athens (Greece), 10-14 July 2016 7