International Symposium on Aircraft Technology, MRO & Operations 2021 June 28-30June 2021 Budapest, Hungary ISATECH-2021-000 Evaluating the Effects of a Morphed Trailing Edge Flap for Aeroacoustic Applications Joseph C. Watkins, Abdessalem Bouferrouk Engineering Design and Mathematics, The University of the West of England, Bristol, BS16 1QY, UK. Joewatkins99@live.co.uk, abdessalem.bouferrouk@uwe.ac.uk SUMMARY This paper sumamrises the results of 2D CFD simulations of a 30P 30N aerofoil section with and without a morphing trailing edge f lap ( TEF). Numerical simulations are carried out at a f ixed angle of attack of 8 -degrees, with f lap deflections of 5 - and 15 - degrees. Results demonstrate that in some cases the morphing TEF is able to reduce noise without reducing aerodynamic efficiency, whilst in other cases the overall performance is similar to that of a f lapped conventional configuration. Keywords: Aeroacoustics, 30P30N high lift wing, CFD, morphing, trailing edge flap. INTRODUCTION The aim of this project is to investigate the effectiveness of morphing as a technology for noise reduction on high lift devices, primarily the noise generated by the gap between a wing and an extended trailing edge flap (TEF). Morphing has been recently studied in many papers (e.g. Kimaru and Bouferrouk, 2017), and this particular paper adds to some previous work on morphing TEFs (Evans et al., 2016, Loudon et al., 2018). Aircraft engine noise has been reduced significantly through innovations such as the mass production of high bypass ratio turbofan engines (e.g., the Rolls Royce Trent engine line), leading to airframe noise becoming a much larger contributor to overall aircraft noise than it used to be. This study uses the 30P30N aerofoil, with no leading-edge slat, to analyze the far field aeroacoustic noise propagation from the flap cove region, and any other characteristic noise sources. Further analysis assesses whether this noise can be reduced by morphing the region linking the main wing and the flap, to create one continuous smooth surface. The leading-edge slat on the 30P30N aerofoil geometry is not included in these experiments as not all aircraft utilize a leading-edge slat; and so, innovations in flap noise reduction techniques can be applied to a wider range of aircraft. Projects such as this have increased in relevance recently because residential areas surrounding airports are becoming increasingly more densely populate. Past research has shown the damaging effects of loud noise near residential areas by disturbing sleep, causing stress, damaging surrounding ecosystems and reducing wildlife populations. The 30P30N aerofoil is widely used for studying the aeroacoustics of high lift devices, along with various reduction techniques. The 30P30N aerofoil geometry is shown below in Figure 1. Figure 1: 30P30N aerofoil. Murayama, M. et al. (2014) tested the 30P30N aerofoil in the 2x2 meter JAXA wind tunnel test section. They placed a series of unsteady pressure transducers on the surface of the wing to analyse the near field sound propagation, and an array of microphones on the walls of the test section to assess the far field noise. They concluded that the 30P30N aerofoil has multiple narrow broadband peaks across all angles of attack tested, which decreased in amplitude as the angle of attack was increased. Similarly, they found that broadband noise for this aerofoil configuration is at its highest when at low angles of attack (i.e., 0 and 3.5 degrees). Using data from the near field pressure transducers around the aerofoil geometry, they demonstrated that the highest contributor to the overall wing section noise originates from the leading-edge slat. More recently, research work by R.Ilário da Silva et al. (2020) investigated the aeroacoustic effects of using various designs of fillers for the cove regions on the slat and the flap of the 30P30N aerofoil. The geometries they test are shown in Figure 2. International Symposium on Aircraft Technology, MRO & Operations 2021ISATECH-2021-000 Operations 2021 June 28-30 2021 Budapest, Hungary aerofoil; and 25 chord lengths behind, to enable the wake flow to develop fully. A C-shaped domain is also used because triangular elements are used, with a sharp trailing edged geometry. This produced a mesh with very good skewness and orthogonal qualities. Mesh Meshing is done using a hybrid approach, with structured inflation layers near the wall. The first layer height (FLH) is calculated using the chord length of 0.45m, and a target y+ of 1. This is to create a mesh with appropriately small near-wall element size to capture the turbulent boundary layer accurately. The total number of layers is calculated using Blasius’ turbulent boundary layer height equation. The FLH used for testing is 6.45 × 10−6 m, with 45 inflation layers, and a geometric growth rate of 1.2. The rest of the domain element size is determined by a wall element size of 9.7 × 10−4 m, and a small element growth rate of 1.03 to an outer domain size of 0.4m. This gradually phases out the acoustic pressure waves in such a way that they are captured accurately by the receiver but are phased out before reaching the domain boundaries. This reduces the impact of the waves reflecting off the boundaries and being detected twice by the receiver. The final mesh and domain can be seen in Figure 4. The same mesh parameters are used across all configurations. The maximum wall y+ across all meshes used is 1.2; and there are around 200,000 elements. Figure 2: Experimental configurations (Silva et al. ,2020). Their results showed that there was no significant change in lift or drag ratio except for the full slat cove filler if used, which resulted in an improved lift to drag ratio. The acoustic results showed that the full cove filler configuration reduced the overall broadband noise across a wide spectrum by 4-5 dB, and they almost eliminated the tonal noise generated by the sharp trailing edge of the slat. More research is still required to isolate the flap noise and create noise reduction techniques for the noise generated by the secondary cove region for the trailing-edge slat. This paper investigates the use of a morphing trailing edge flap for noise abatement. METHODOLOGY Geometry This study uses ANSYS Fluent, a cell centered Computational Fluid Dynamics (CFD) software, using a 2D planar geometry. The effects of fully morphing the trailing edge flap of the 30P30N aerofoil to the main body are investigated. This is done using a series of Bsplines to produce a smooth continuous surface from the trailing edge of the main body to the leading edge of the flap. Testing is carried out at 8 degrees angle of attack, with two different flap deflections at 5- and 15degrees. The conventional and the morphed configurations are shown in Figure 3 at 15-degree flap deflection. The same technique is also applied to the 5-degree case. Figure 4: Domain mesh and geometry configuration for conventional flap at 15-degrees flap deflection Solver Method The computational acoustic analogy used consists of firstly allowing the simulation to run as steady for around 1000 iterations; this acts as a good initialization for the transient simulation to speed up convergence. Following this, the simulation is changed to transient, with a time step calculated using Equation 1, where the Courant number is 1, flow velocity is 30m/s and Δπ₯ is the smallest mesh element size. This results in a time step of around 5.8 × 10−5 seconds (varying slightly for different configurations due to small mesh differences). Figure 3: 15-degree flap deflection for conventional (above) and morphed (below) configurations The computational domain boundaries are located at 13 chord lengths in front, above and below the 2 International Symposium on Aircraft Technology, MRO & Operations 2021ISATECH-2021-000 Operations 2021 June 28-30 2021 Budapest, Hungary ππππ π π‘ππ (Δπ‘) = 80.00 70.00 60.00 50.00 40.00 30.00 πΆππ’ππππ‘ ππ’ππππ × π∞ Δπ₯ Equation 1 20.00 10.00 0.00 -10.000 -20.00 π πππ .00 5000 .00 10000.00 -30.00 Frequency (Hz) Figure 5: SPL Vs Frequency plot for morphed and conventional configurations at 5-degree flap deflection The large main feature is the initial low frequency tone which peaks for both configurations at around 250Hz, with a similar amplitude of 60 dB. This wide frequency band peak is caused by the vortex shedding from the trailing edge of both aerofoils. The shedding vortex is large and rotates at a lower frequency, creating a loud low frequency noise. Both aerofoils low frequency noise characteristics are quite similar due to them having the same chord length and are at the same angle of attack. After 1000Hz, however, the morphed configuration is much quieter. On average, the broadband noise generation of the morphed configuration across the wide spectrum of frequencies, is around 25% quieter than the conventional. This increased broadband noise generation is likely due to increased flow separation caused by the cove region on the conventional flap design. The flow separation region often has smaller eddies which can rotate across higher frequencies, generating a broadband type of noise. There are also two distinct tonal peaks produced by the conventional configuration at 2400Hz and 4800Hz, with an amplitude of 62 dB and 50 dB respectively. These tonal peaks are not present in the morphed configuration. The expected cause for these tonal peaks in the conventional flap design, is likely to be the additional sharp trailing edge of the main body of the 30P30N geometry, e.g. Peng et al. (2018). This results in a sharp release of turbulent kinetic energy from the trailing edge from the cove region and generating tonal noise. the PRESTO! solution method is used for pressure to achieve greater accuracy for pressure fluctuations. The simulation run time with acoustics model activated, is determined using Equation 2; this is based on running 5 periods for the corresponding minimum obtained frequency. (ππππ − ππππππ ) = π Conventional SPL (dB) The simulation is then run until a statistically steady state has been achieved (this is defined as the flow having no overall uptrend or downtrend in specific parameters such as lift and drag coefficient). Next, data sampling for time statistics is turned on to obtain comparable pressure and aerodynamic parameters averaged over time. In addition to activating the Ffowcs-Williams and Hawkings (FWH) acoustic model with a reference pressure of 2× 10−5 Pa (the lower threshold of human hearing), the source correlation length is set to 0.53m, which was also used by Jawahar et al. (2021) whose wind tunnel acoustic data is used for validation of the acoustic analogy. The receiver is located at 1.5m below the leading edge of the aerofoil (around 3 chord lengths as to be defined as far field). The Coupled pressure-velocity coupling scheme is used for all unsteady simulations, along with a second order upwind scheme for all flow variables. However, Morphed Equation 2 RESULTS AND DISCUSSION At 5-degrees flap deflection, the aerodynamic performance of the morphed configuration is slightly better than the conventional, with 7% more lift generated, and negligible differences in drag. This results in lift to drag ratios of 18 and 16 for the morphed and the conventional configurations, respectively. When the flap deflection is increased to 15degrees, the differences in aerodynamic performance between the two configurations are very small. However, the aerodynamic performance of the conventional flap configuration (with the gap) improves drastically as the flap deflection is increased, whereas the morphed flap’s performance improving only slightly. The morphed configuration sees a much greater increase in drag over the conventional configuration as the flap deflection is increased. This is likely due to the maintained higher-pressure difference between the upper and lower surface in the morphed flap case, which results in a more turbulent wake flow from the trailing edge. As the flap deflection is increased to 15 degrees, the disadvantages observed in the conventional design decrease, and the two configurations havemore similar noise characteristics across the frequencies tested. Figure 6 shows the SPL generationwith respect to frequency at 15-degree flap deflection for both configurations. The acoustic noise generation of both configurations at 5-degree flap deflection is shown in Figure 5 in a sound pressure level (SPL) versus frequency plot. SPL plots are created by carrying out a Fast Fourier Transform (FFT) on pressure data. 3 International Symposium on Aircraft Technology, MRO & Operations 2021ISATECH-2021-000 Operations 2021 June 28-30 2021 Budapest, Hungary ACKNOWLEDGEMENT 80.00 Morphed conventional The authors are grateful for the support of the University of the West of England’s IT services to run the CFD simulations. 60.00 40.00 SPL (dB) NOMENCALTURE CFD FWH FFT π‘πππ π‘π π‘πππ‘ Δπ‘ ππππ π∞ Δπ₯ 20.00 0.00 0. 00 5000.00 10000.00 -20.00 -40.00 Frequency (Hz) Figure 6: SPL Vs Frequency plot for morphed and conventional configurations at 15-degree flap deflection Computational Fluid Dynamics Ffowcs-Williams and Hawkings method Fast Fourier Transform Flow time at end of simulation (seconds) Flow time at beginning of simulation (seconds) Numerical time step (seconds) Minimum frequency obtained in simulation (Hz) Free stream velocity (m/s) Smallest element size in flow direction (m) REFERENCES Kimaru, J. Bouferrouk, A. (2017) Design, manufacture and test of a camber morphing wing using MFC actuated smart rib. 8th International Conference on Mechanical & Aerospace Engineering, Prague, 22- 25 July 2017. Prague: IEEE, pp. 791796. Like the 5-degree deflection results, there is a distinctly loud, low frequency peak at 350Hz for both configurations, peaking slightly higher than at 5degrees, up to 70dB for both. Again, this is likely caused by the large eddy vortex shedding in the wake. The broadband noise generation of the conventional design is now nearly the same level as the morphed configuration, which begins to be even quieter at the highest frequencies. The two distinct tonal peaks are still mildly present in the conventional design at 15-degrees but have a much smaller amplitude. This effect is like what Murayama, M. et al. (2014) reported in that as the angle of attack of the 30P30N aerofoil increases, the narrow broadband peaks are significantly reduced. Evans, C., Harmer, M., Marks, O., Tiley, S., Willis, T., Bouferrouk, A. and Yao, Y. (2016) Development and testing of a variable camber morphing wing mechanism. In: International Symposium of Sustainable Aviation (ISSA), Istanbul, Turkey, 29 May - 1 June 2016. Loudon, K., Bouferrouk, A., Coleman, B., Hughes, F., Lewis, B., Parsons, B., Cole, A. and Yao, Y. (2018) Further development of a variable camber morphing mechanism using the direct control airfoil geometry concept. In: International Symposium of Sustainable Aviation, Rome, Italy, 9-11 July 2018. Murayama, M., Nakakita, K., Yamamoto, K., Ura, H., Ito, Y. and Choudhari, M., (2014). Experimental Study on Slat Noise from 30P30N Three-Element High-Lift Airfoil at JAXA Hard-Wall Lowspeed Wind Tunnel. CONCLUSION R.Ilário da Silva, C., Azarpeyvand, M., Alihan Showkat, S. and Kamliya Jawahar, H., (2020). Aerodynamic and aeroacoustic performance of high-lift airfoil fitted with slat cove fillers. The research presented in this paper illustrates how morphed trailing edge flaps can affect the acoustic noise as well as aerodynamic performance when compared with conventional trailing edge flaps. Using an FFT of pressure data to produce SPL plots, results show that the morphed configuration holds significant benefit over the conventional configuration at low flap deflections; where there is 25% less broadband noise, and no tonal peaks when compared with the conventional design. In addition to aeroacoustic benefits, the morphed configuration sees a 12% increase in lift to drag ratio over the conventional. These advantages of the morphed configuration decrease when the flap deflection is increased. The distinct tonal peaks in the conventional gapped flap design reduce significantly, and broadband noise generation is decreased to nearly the same level as the morphed configuration. These results demonstrate promising attributes of a fully morphed trailing edge flap, which should be investigated further. Key future work includes testing a wider range of flap deflections and angles of attack; experimental wind tunnel testing; and whether a practical mechanical mechanism can be developed so that if the technology is proven to be effective, it can be applied to full scale aircraft. Peng, T., Yao, Y. and Zhu, Q., (2018). Slat broadband noise prediction of multi-element 30P30N airfoil by a hybrid RANS-LES method. 4