This work presents comparisons between experimental and numerical estimates of near-field rotor

aeroacoustics in hover. The experiments took place at the Kazan National Research Technical University

named after A. N. Tupolev (Kazan Aviation Institute). A set of rotor blades with NACA-0012 aerofoil sections

was used to obtain the sound pressure distribution using a linear array of microphones. It is shown that CFD

and experimental results are in good agreement suggesting that the obtained test data can be useful as a

validation case for development of CFD tools

Barakos, George N.

Batrakov, Andrey

Bozhenko, Andrey

Kusyumov, Alexander

Mikhailov, Sergey

Pakhov, Vladimir

Stepanov, Robert

English

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Paper #  Page 1 of 6  Presented at 44th European Rotorcraft Forum, Delft, The Netherlands, 19-20 September, 2018  This work is licensed under the Creative Commons Attribution International License (CC BY). Copyright © 2018 by author(s).  EXPERIMENTAL AND NUMERICAL INVESTIGATION OF NEAR-FIELD ROTOR AEROACOUSTICS Robert Stepanov, rpstepanov@kai.ru,        Kazan National Research Technical University named after A.N. Tupolev,      (Russian Federation) Andrey Batrakov, asbatrakov@kai.ru,         Kazan National Research Technical University named after A.N. Tupolev       (Russian Federation)    Vladimir Pakhov, VVPakhov@kai.ru,         Kazan National Research Technical University named after A.N. Tupolev       (Russian Federation) Alexander Kusyumov, ANKusyumov@kai.ru,    Kazan National Research Technical University named after A.N. Tupolev       (Russian Federation) George N. Barakos, George.Barakos@glasgow.ac.uk, University of Glasgow   (United Kingdom) Andrey Bozhenko, anbozhenko@kai.ru,        Kazan National Research Technical University named after A.N. Tupolev        (Russian Federation) Sergey Mikhailov, sergey.mikhaylov@kai.ru, Kazan National Research Technical University named after A.N. Tupolev       (Russian Federation)    Abstract This work presents comparisons between experimental and numerical estimates of near-field rotor aeroacoustics in hover. The experiments took place at the Kazan National Research Technical University named after A. N. Tupolev (Kazan Aviation Institute). A set of rotor blades with NACA-0012 aerofoil sections was used to obtain the sound  pressure distribution using a linear array of microphones. It is shown that CFD and experimental results are in good agreement suggesting that the obtained test data can be useful as a validation case for development of CFD tools.  1. INTRODUCTION1 In recent years, the noise reduction of helicopters has become a priority due to new, stricter certification requirements for civil helicopters, operating in densely populated areas. Obtaining reliable experimental data of rotor aeroacoustics is a challenging task due to the high cost of experimental facilities, and the need for wind tunnels with high operational costs. In recent years, attention has been shifting towards numerical simulation, where experimental data is needed to validate the obtained numerical                                                        Copyright Statement The authors confirm that they, and/or their company or organization, hold copyright on all of the original material included in this paper. The authors also confirm that they have obtained permission, from the copyright holder of any third party material included in this paper, to publish it as part of their paper. The authors confirm that they give permission, or have obtained permission from the copyright holder of this paper, for the publication and distribution of this paper as part of the ERF proceedings or as individual offprints from the proceedings and for inclusion in a freely accessible web-based repository.  results [1-3].  This work aims to obtain experimental data in near-field, suitable for validation of CFD computations. To this end, a set of blades of a rectangular planform were manufactured using the NACA-0012 aerofoil. A linear microphone array was placed next to the Mach-scaled rotor model, and hence the measurements were dominated primarily by thickness noise, which is dictated mainly by the angular speed of the rotor and blade thickness. The thickness noise is the dominant component on the rotor plane, especially in hover [4].   The broadband noise sources are related to turbulence and are caused by the interaction between blades and turbulence; self noise of the boundary layer of the blade [5, 6], and turbulence behind its sharp trailing edge. The mechanisms of rotor broadband noise generation are described in [7]. An advantage of the current dataset is that near-field acoustic pressure is made available to allow direct comparisons with CFD computations.  CFD computations were obtained using the HMB code of Glasgow University and were based on Page 2 of 6  Presented at 44th European Rotorcraft Forum, Delft, The Netherlands, 19-20 September, 2018  This work is licensed under the Creative Commons Attribution International License (CC BY). Copyright © 2018 by author(s). RANS approach with the k-ω turbulence model. The numerical results were then compared to the experimental results as sound pressure level (SPL) in the time domain. Previous studies for a different set of blades showed good agreement of experimental data with CFD computations [2, 3].   2. EXPERIMENTAL SETUP AND  CONDITIONS  2.1. Wind Tunnel Experiments were conducted in the acoustic chamber of T-1K wind tunnel of the Kazan National Technical University named after A.N. Tupolev. The acoustic chamber consists of retractable side walls, which contain Helmholz resonators to minimize low frequency noise in the test section, and melamine pyramid-shaped foam material to absorb high-frequency noise, emanating from rotors, operating at high tip speeds. The acoustic chamber is shown in   Figure 1.  Figure 1. 3D model of the acoustic chamber of T-1K wind tunnel at KNRTU-KAI.  2.2. Rotor Rig  All experiments were performed in hover. The rotor rig was operated at 900 rpm, corresponding to tip Mach number of 0.22. The rotor radius was R=0.8 m. The collective pitch angle was set to 8°. The rotor blades with a rectangular planform had a constant cord with a symmetrical NACA-0012 aerofoil, and no twist along their span.  Experimental results are presented in terms of relative radius ?̅? = 𝑟/𝑅, where 𝑟 is a horizontal distance from the rotor's axis of rotation; and 𝑦ത =𝑦/𝑅, where 𝑦 is the vertical distance of a microphone from the rotor plane.  2.3. Data Acquisition and Analysis For the recordings, DMX RTA-M microphones with Panasonic WM-61A cartridges were used. A 24-bit NI-PXI 4496 ADC sampled the signals at a rate of 48 kHz.  Experimental data was obtained using a linear array of 9 microphones. The placement of the linear array relative to the rotor is shown in Figure 2. The array was positioned vertically at a distance of ?̅? = 1.23 away from the rotor axis of rotation. Earlier work showed that positioning microphones in this manner has little effect on the flow field of the rotor [3].  Figure 2. Placement of microphones relative to the rotor. Due to the periodicity of the microphone signal with respect to the  angular position of the blade, phase-averaging was used, calculated as follows: (1)                       𝑝௔௩௚ =ଵே∑ 𝑝௜௝ே௜ୀଵ . Here 𝑖 is the period of the passing blade, and 𝑗 is a pressure reading 𝑝௜௝  along the period 𝑖. The phase averaged confidence intervals 𝜎௝ were calculated over 60 periods as the root mean square deviation: (2)                  𝜎௝ = ටଵே∑ ൫𝑝௔௩௚ − 𝑝௜௝൯ଶே௜ୀଵ . Compared to the results, presented in the earlier work for a different set of blades [3], the experimental data for each microphone position is analyzed over a larger number of periods (𝑖 = 60), which could contribute to a wider range of confidence intervals of the experimental data, obtained in this work. Page 3 of 6  Presented at 44th European Rotorcraft Forum, Delft, The Netherlands, 19-20 September, 2018  This work is licensed under the Creative Commons Attribution International License (CC BY). Copyright © 2018 by author(s). 2.4. Some Acoustic Characteristics of the  Anechoic Chamber A preliminary study has been performed to investigate noise reflections inside the anechoic chamber.  For this purpose, the Maximal Length Sequence (MLS) [8] method was used. It requires a noise source with volumetric velocity measuring, which is based on a given noise signal with its autocorrelation function close to a delta-function. Figure 3 shows the schematic of the experimental setup. Figure 3. Schematic of the reflection experiment   Figure 4 shows measurement results of the reflection study for microphone #6. Here, the reflection peaks are shown as a function of distance from the microphone. The impulse at 2,5 m corresponds to reflections from side walls, and peaks at 4,5 m correspond to reflections from the floor. It can be seen that the noise reflections are 6-6.5 times lower than the initial impulse for both cases. Figure 4. Reflection study results for    microphone 6.  3. COMPUTATIONAL SETTINGS CFD simulations based on RANS solutions with the k-ω turbulence model and were performed using the HMB code [9]. Due to the periodicity of the flow around a rotor operating in hover mode, the computational domain was set up only for one blade, as shown in Figure 1. A multiblock grid was built using ICEM HexaTM. The grid contained 4.4 million points and 172 blocks. The mesh was put together based on experience from previous studies, and was designed to give a high resolution flow field with good accuracy [10]. The wake topology from CFD simulations is shown in Figure 6. The ground proximity of the experimental setup was not taken into account during CFD simulations, due to the expected small influence on the results.   Figure 5. Computational domain.   Figure 6. Wake topology obtained from CFD simulations. Paper #  Page 4 of 6  Presented at 44th European Rotorcraft Forum, Delft, The Netherlands, 19-20 September, 2018  This work is licensed under the Creative Commons Attribution International License (CC BY). Copyright © 2018 by author(s). 4. RESULTS AND DISCUSSION Comparison of CFD with experimental results for the SPL distribution, obtained from the linear array, is shown in Figure 7. Here, SPL values are presented in terms of peak-to-peak decibel values of the signal. The acoustic noise peaks in the vicinity of the rotor plane, reaching its highest slightly above the rotor blade. It should be noted that the CFD models had no blade coning angle. However, the rotor rig model had a finite coning angle due to the flap hinge of the rotor head. It can be seen from Figure 7 that SPL peaks at 𝑦ത = 0.163 for both CFD and the experiment, which indicates that it cannot be solely attributed to the coning angle.  It can also be seen that the CFD simulations agree well with the experimental data. A noticeable difference of 7 dB can be seen between experiments and CFD at  𝑦ത = −0.65, which happens due to the very low values of peak-to-peak pressure amplitudes in that region. Figure 8 shows this clearly.  Figure 7. Acoustic pressure distribution of the linear array, located at a distance ?̅? = 1.23.  Figure 8 shows comparisons of CFD and experimental results for the acoustic pressure distributions in the time domain. Statistically analyzed experimental data is presented along with their confidence intervals, as mentioned in Section 2.3. Results of CFD simulations are denoted by long dash lines. It can be seen that CFD and experimental data are in good agreement, and most of the time, pressure distributions obtained from CFD, lie within the phase averaged confidence intervals. However, a noticeable difference of the signal can be observed after the pressure peaks for microphone positions 𝑦ത = 0, 𝑦ത = 0.163 and 𝑦ത = 0.325, where CFD has lower pressure levels compared to the experiment. It should be stated that these microphone locations correspond to highest peaks of the SPL in Figure 7. Experimental SPLs also show somehow larger values compared to CFD below the rotor plane, at microphone positions 𝑦ത = −0.163 and 𝑦ത = −0.325, although the overall signal amplitude tends to be smaller below the rotor plane. The slightly higher CFD estimates at 𝑦ത = 0.163 ahead of the blade are similar to results obtained for a different set of blades [3].   6. CONCLUSIONS Experimental results and CFD simulations for near-field rotor acoustics were presented and compared.  A set of blades with a rectangular planform and NACA-0012 aerofoil have been used, and data was gathered using a linear array of microphones. The agreement between simulations and experiments is, overall, very good, suggesting that this data set can be used as a first step in the validation of CFD codes for near-field acoustic predictions. This study is expected to move towards measurements of different blade shapes and also cover rotors in forward flight.  7. NOTATION 𝑟 Horizontal distance from the rotor's axis of  rotation.  ?̅? Relative distance, scaled to the rotor  radius (?̅? = 𝑟/𝑅). 𝑦 Vertical distance from the rotor plane.    𝑦 is positive in the upstream direction. 𝑦ത Relative vertical distnace, scaled to the  radius 𝑅 (𝑦ത = 𝑦/𝑅 m). 𝑅 Rotor radius (𝑅 = 0.8 m). SPL Sound pressure level.  8. ACKNOWLEDGEMENTS This work was supported by the grant ‘Numerical and physical modelling of aerodynamic and aeroacoustic characteristics of rotor systems of future concept aircarft’ (No. 9.1577.2017/4.6) of the Ministry of Education and Science of the Russian Federation.  9. REFERENCES [1] Kube, R., Splettstoesser, W.R., Wagner, W., Seelhorst, U., Yu, Y.H., Tung, C., Beaumier, P., Prieur, J., Rahier, G., Spiegel, P., Brooks, T.F., Burley, C.L., Boyd, D.D., Mercker, E., Pengel, K, "HHC Aeroacoustic Rotor Tests in the German-Dutch Wind Tunnel: Improving Physical Paper #  Page 5 of 6  Presented at 44th European Rotorcraft Forum, Delft, The Netherlands, 19-20 September, 2018  This work is licensed under the Creative Commons Attribution International License (CC BY). Copyright © 2018 by author(s).   Figure 8. Comparison of experimental data with CFD results for different vertical distances 𝑦ത of the microphones.  Paper #  Page 6 of 6  Presented at 44th European Rotorcraft Forum, Delft, The Netherlands, 19-20 September, 2018  This work is licensed under the Creative Commons Attribution International License (CC BY). Copyright © 2018 by author(s). Understanding and Prediction Code." Aerospace Science and Technology, 1998, no. 3, 177-190. [2] Pakhov, V., Stepanov, R., Bozhenko, A., Batrakov, A. , Kusyumov, A. Mikhailov, S.   Barakos, G.N., "Creating a database of helicopter main rotor acoustics for validation of CFD methods." 7th European Conference for Aeronautics and Space Sciences (EUCASS), Milan, Italy, paper 638 (2017). [3] Stepanov, R., Pakhov, V., Bozhenko, A., Batrakov, A., Garipova, L., Kusyumov, A., Mikhailov, S., Barakos, G.N, "Experimental and numerical study of rotor aeroacoustics." International Journal of Aeroacoustics. Vol 16, Issue 6, pp. 460–475, 2017. [4] Brentner, K.S., Farassat, F., "Modeling aerodynamically generated sound of helicopter rotors." Progress in Aerospace Sciences, 39, pp. 83 – 120, 2003. [5] Farassat, F. and Farris, M., "Verifcation and analysis of 4 formulation of langley for the study of noise from high speed surfaces". Proceedings of the 5th Joint CEAS / AIAA Aeroacoustics Conference, Vol. I, 1999. CEAS/AIAA Paper     99-1881. [6] Glegg, S.A.L. and Jochault, C. "Broadband self-noise from a ducted fan". AIAA Journal, No. 36, vol.8, pp. 1387-1395, 1998. [7] Burley, C.L., and Brooks, T.F. “Rotor Broadband Noise Prediction with Comparison to Model Data.” Journal of American Helicopter Society, 49:1, January 2004. [8] Ostrikov, N., Denisov, S., Makashov, S., Anikin, V., Gromov, V. “About estimating of acoustic characteristics of helicopter`s main rotor models at opened stand conditions”// Acoustic Journal, 2016, Vol.62, Issue 6, pp. 725-730 (In Russian) [9] Steijl R and Barakos GN. Sliding mesh algorithm for CFD analysis of helicopter rotor–fuselage aerodynamics. Int J Numer Methods Fluids 2008; 58: 527–549. [10] Garipova LI, Batrakov AS, Kusyumov AN, et al. Aerodynamic and acoustic analysis of main rotor blade tip part for hover. Int J Numer Methods Heat Fluid Flow 2016; 26: 2101–2118.     

Experimental and Numerical Investigation of Near-field Rotor Aeroacoustics

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Experimental and Numerical Investigation of Near-field Rotor Aeroacoustics

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