The geometric configuration used for validation was the default 30P30N aerofoil (three elements at 30° flap and slat deflection). All experimental conditions of the wind tunnel experiments, replicated in Ansys Fluent—with a Reynolds number Re =
(Mach number M = 0.17)—were used for pressure and aerodynamic performance validation, whereas conditions of Reynolds number Re =
(Mach number of M = 0.089) were used for aeroacoustic validation. Firstly, pressure plots from Murayama et al. [3
] were used to validate the CFD methodology, as shown in Figure 9
for both 3° and 8° angles of attack (AoAs). The X-location in Figure 9
was measured starting from the slat’s leading edge and ending at the flap’s trailing edge.
The segments of the plots show the slat, main body, and flap pressure coefficient distributions going from left to right. At the leading edge of each component, there is a large pressure difference between the upper and lower surfaces, where a large amount of the lift is generated. The pressure coefficient plots show good correlation with the wind tunnel data at both angles of attack, but slightly overpredict the pressure across all surface locations. Overall, the numerical data show nearly the same trend as the wind tunnel data, with exceptionally close amplitudes. Following this, the lift coefficient was plotted against both the wind tunnel data and the CFD data of Murayama et al. [3
], as seen in Figure 10
The data were nearly the same up to 18°, with an error of only around 5%. Murayama et al. [3
] did not show experimental lift coefficient data beyond 18°; however, they stated that they found that their CFD data predicted stall at a much higher angle of attack than the wind tunnel. The current results follow the same trend as the CFD data of Murayama et al. [3
], and show good correlation with their experimental data.
Finally, the acoustic simulation methodology was validated against the wind tunnel experiments of Jawahar et al. [4
]. Results are shown in Figure 11
for the sound pressure level (SPL) in dB versus Strouhal number, defined as:
is the vortex shedding frequency, c
is the chord, and
is the freestream velocity.
There was generally a reasonable agreement between the two datasets in the SPL noise levels across the range of Strouhal numbers shown, from 0.2 to 10. Lower Strouhal numbers were mostly the same, with the CFD data predicting noise to be slightly higher than wind tunnel data. Peaks across all Strouhal numbers occurred later in the CFD data, with the wind tunnel tonal peaks at Strouhal numbers of 1.6 and 3.2 being much smaller and at a higher frequency compared with the numerical data. The current CFD simulation did not fully resolve all tonal peaks, and primarily predicted the incorrect amplitude of the wind tunnel peaks at Strouhal numbers of 1.6 and 3.2. Due to a variety of measurement uncertainties that may be encountered in a wind tunnel, validating numerical acoustic data against experiments is always challenging. Despite this, the numerical methodology overall produces reasonably accurate results, including the general trend, compared with those from the wind tunnel tests.
3.2. Results of Morphed and Conventional Flap Design
Results for both conventional and morphed flap designs are now presented. The focus was on the far-field acoustic noise generated by the fluid interaction with the high-lift aerofoil, along with aerodynamic performance analysis. All simulations were carried out at an 8° angle of attack and a Reynolds number of (the same conditions as in the validation study, except there was no leading-edge slat). When calculating the time-averaged and unsteady flow characteristics, the initial transients were disregarded.
Although the meshes for all cases were kept more or less the same, there were bound to be small changes in the smallest mesh element size—especially with different flap deflections—and with the condition to keep the Courant number at 1, there were slight changes in the timestep used (due to the smallest mesh element being in the timestep calculation); hence, there were very slight variations in the maximum frequency resolved (according to the Nyquist criterion). For the SPL results in Figure 12
, this slight change in the upper frequency has no bearing on the noise level at high frequencies, and the reason the plots are shown for the same upper frequency cutoff is purely for convenience. For all computations of SPL in Figure 12
, a C
value of 1 was used.
The main low-frequency acoustic feature, which is present in both the morphed and conventional aerofoil designs, can be seen in Figure 12
between 0 and 1 kHz frequency at all flap deflections. Generally, low-frequency noise is caused by larger eddies, as they require much more energy to rotate at a higher frequency. This low-frequency peak may be associated with regular vortex shedding from the trailing edge of the flap. It can be seen that the level of noise in the low-frequency peak of the morphed configuration increases slightly with the flap deflection angle, presumably due to increased interaction of the main wing’s flow with the flap. However, for the conventional design, the low-frequency peak remains almost the same for all flap deflections. Both flaps have the same chord length and operate at the same flap deflection and flow velocity, and will therefore have similar-sized vortex shedding at their trailing edge.
Across all five flap deflection angles tested in the conventional flap design case, there were three large tonal peaks, which occurred at frequencies of 2.5 kHz, 5 kHz, and 7.5 kHz (apart from the 5° deflection, where there were only two tonal peaks at 2.5 kHz and 5 kHz). These tonal peaks increased in amplitude as the flap deflection was increased. The conventional flap’s tonal peak amplitude increase with flap deflection is possibly caused by the increased pressure difference between the upper and lower surfaces at higher flap deflections which, when the gap shear layer and wing trailing-edge boundary layer mix, create a much louder tonal peak as the combined flow sheds downstream [11
]. These tonal peaks are not present at low deflection angles in the morphed case, but do form at 20 and 25° flap deflection, although they are still relatively small.
The actual cause of these small tonal peaks that form in the morphed case at higher flap deflections is unclear, and is a little surprising given that the contoured flap surface is connected smoothly to the outer main wing. It seems that the morphed flap causes surface pressure fluctuations at its trailing edge that are higher than those generated at the trailing edge of the conventional flap. A similar observation was reported by Jawahar et al. [18
] for their comparison of the noise generated by a conventional hinged flap and by a morphed flap of a NACA 0012 aerofoil.
In addition to the absence of mid-to-high-frequency tonal peaks, the morphed configuration has a much lower level of broadband noise across all flap deflections tested. One aerodynamic mechanism that is usually associated with broadband noise generation is flow separation. The conventional flap design, due to its cove region, experiences much larger areas of flow separation than the morphed configuration. It is expected that the flow separation in these cove regions is the cause of the higher levels of broadband noise seen in Figure 12
in the conventional flap.
The broadband noise for both the conventional and morphed flap designs can be more easily compared when looking at the 3rd octave band plots, as shown in Figure 13
(conventional flap) and Figure 14
In Figure 13
, it can be observed that as the flap deflection increases, the level of broadband noise and the amplitude of the tonal peaks increases; this is characterised by the higher levels of flow separation and flow unsteadiness caused by the unsteady mixing shear layer from the main body’s trailing edge. The 3rd octave band plot in the morphed configuration (Figure 14
) does not have such a straightforward and obvious trend. It is not clear which flap deflection generates the least levels of broadband noise, as the sound pressure level at any given frequency range varies so much with flap deflection. However, when a mean sound pressure level is computed for each frequency band across all flap deflections, it gives a better picture of the overall noise generation, and it can be seen that the morphed configuration still has a significantly lower broadband noise than the conventional one. The plot in Figure 15
shows an average of 11% less noise across all frequencies obtained, and 20% less broadband noise when considering only the frequency bands above 1250–1600 Hz.
Overall, the morphed-flap design clearly holds the most preferred characteristics when it comes to the aeroacoustic performance, with an average of 11% less broadband noise, and the elimination of tonal peaks present in the conventional case. However, as Whitfield [2
] stated in 2004, after noise-generation mechanisms have been identified and noise-reduction techniques developed, the key is to implement them without having to compromise extensively on aerodynamic performance.
The aerodynamic characteristics of lift and drag coefficients (
) and their combined ratio
are shown in Table 1
and Table 2
for the conventional and morphed flap configurations, respectively, for the range of flap deflections studied.
Whilst the conventional flap stalls at around 20°, the morphed flap does not stall until after 25° flap deflection. Evidence of the delayed stall in the morphed flap is supported by Figure 16
for the case of 5
flap deflection, where the flow seems to be more effectively attached to the upper surface towards the trailing edge of the flap than in the conventional case. The flow’s ability to remain attached for longer on the morphed configuration is due to the continuous surface, leading to smooth flow. Flap stall occurring earlier in the conventional flap design is due to increased flow separation, which impairs the aerofoil’s ability to maintain low pressure on the upper surface, leading to decreased lift generation.
In terms of the lift-to-drag ratio, the morphed flap exhibits a superior aerodynamic performance at flap deflections below 15°; after this flap deflection, however, and due to the rapid increase in pressure drag relative to the conventional flap, the conventional flap design has a higher lift-to-drag ratio.
In the conventional flap design, some pressure difference between the upper and lower surfaces is equalised at the cove region, as depicted by the relatively high levels of turbulent kinetic energy (TKE) in the cove region of the conventional flap in Figure 16
. Therefore, there is less turbulence and overall energy in the flow as the upper and lower surface shear layers mix at the trailing edge to form the unsteady downstream flow. In the morphed case, however, all of the pressure difference and turbulent kinetic energy is released at the flap’s trailing edge as the shear layers mix, creating a much more turbulent region after the trailing edge—especially at high flap deflections. The relatively large amounts of TKE immediately after the trailing edge of both configurations are also due to the 30P30N flap having a squared trailing edge (with a thickness); this feature is commonly known to be detrimental for aerodynamics, and has been studied thoroughly in past literature. For example, Thomareis and Papadakis [19
] stated that blunt trailing edges can cause periodic vortex shedding, and can create a spanwise pressure gradient, which is responsible for the development of a secondary flow pattern in the spanwise direction.
To discuss some differences in flow physics between the morphed and the conventional designs, and in support of the initial quantitative results obtained, Figure 17
and Figure 18
are presented for the case of 20° flap deflection. The static pressure contours are plotted in Figure 17
for both the morphed and conventional gapped configurations. It is evident that there is some form of vortex-shedding pattern appearing downstream of the morphing TEF, which is not seen in the conventional setup. It seems that the shear layer formed off the trailing edge of the main wing (as a result of the wing’s TE boundary layer interacting with the shear layer coming from the gap) does not interact with the shear layer forming at the flap’s trailing edge. The two layers flow separately, and only interact and mix further downstream in the wake, by which time the energy dissipated in both flows is not sufficient to cause vortex shedding. One can observe that the shear layer from the main wing’s trailing edge is initially pushed upwards above the flap—presumably due to the high-energy flow from the gap—before it flows downstream almost undisturbed and parallel to the flap. In contrast, for the morphing design, the trailing edge boundary layer flow from the main wing interacts with the flow near the flap’s trailing edge, creating a vortex roll-up and subsequent shedding in the wake. In such a case of 20° flap deflection, one would expect higher pressure drag and lower lift for the morphed configuration (i.e., lower Cl/Cd
ratio, see Table 2
), compared with a higher Cl/Cd
ratio for the conventional reference design (see Table 1
shows the flows streamlines (with velocity magnitudes) near the morphed trailing edge and the trailing edge flap gap for the reference configuration for the case of 20° flap deflection. For the gapped configuration, the streamlines clearly show a large, low-speed, recirculating flow bubble occupying the majority of the gap region, except in the gap’s corner, where there is a much smaller, secondary, recirculating vortex flow. In the simulations, such a recirculation zone seemed to remain steady at various flap deflections, with the secondary zone near the gap corner reducing in size with increasing flap deflection. The flow from the gap interacts smoothly with the main wing’s trailing edge flow, creating a stream of flow that smoothly convects downstream above the flap. Figure 18
clearly illustrates the speeding effect of the flow in the gap—particularly nearer to the flap’s leading-edge region. It is this speed-up effect that helps energise the flap’s boundary layer and keep it attached over most of its upper surface. It can be seen in the simulations that with higher flap deflection angles, the boundary layer on the flap near the trailing edge becomes thicker, and the first signs of vortex sheet roll-up start to appear. For the morphed configuration, the gap flow is eliminated as expected, and with it the speed-up effect over the flap’s leading edge. However, the flap experiences vortex roll-up and shedding, as explained previously from the results of Figure 17