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 Δ𝑥 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
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International Symposium on Aircraft Technology, MRO & Operations 2021ISATECH-2021-000
Operations 2021
June 28-30 2021
Budapest, Hungary
𝑇𝑖𝑚𝑒 𝑠𝑡𝑒𝑝 (Δ𝑡) =
80.00
70.00
60.00
50.00
40.00
30.00
𝐶𝑜𝑢𝑟𝑎𝑛𝑡 𝑛𝑢𝑚𝑏𝑒𝑟 × 𝑈∞
Δ𝑥
Equation 1
20.00
10.00
0.00
-10.000
-20.00
𝟓
𝒎𝒊𝒏
.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.
(𝒕𝒆𝒏𝒅 − 𝒕𝒔𝒕𝒂𝒓𝒕 ) =
𝒇
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.
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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
𝑡𝑒𝑛𝑑
𝑡𝑠𝑡𝑎𝑟𝑡
Δ𝑡
𝑓𝑚𝑖𝑛
𝑈∞
Δ𝑥
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.
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