Flow Visualization

Figure 5.14: Streamlines of the flow in the upper surface of the wing for AoA: 16^◦.

Region Area (m²) Percentage of flow separation (%)

Wing Upper Surface 5.543 -

1 0.4546 8.2

2 0.6588 11.89

3 1.5823 28.55

Total flow separation 48.64

Figure 5.15: Percentage of the boundary flow separation of the upper surface of the wing for an AoA=

16^◦.

In addition, the wing is unswept and due to the interaction between the shear layer that is gene- rated close to the separation line and the wake that roll up in opposite directions, mushroom shaped cells appear just beyond stall, as explained in section 2.3.2. These flow patterns were observed in the streamline contour of the AoA = 18^◦ for three mesh densities: coarse, medium and fine, shown in figure 5.16. The cells are dependent on the aspect ratio of the wing, for the Slingsby aspect ratio of 9, about three cells should appear. In the coarse and medium mesh only a stall cell was observed, as the aircraft is only modelled half, in total there a two stall cells. The fine mesh presents in total four stall cells. However, the empirical correlation between the wing’s aspect ratio and the number of stall cells do not account for the interaction with all the other components of the aircraft and were only performed in rectangular wings. This might explain the different numbers of stall cells observed from the CFD results presented.

Moreover, the negative pressure coefficient as a function of the normalized wing chord was ob- tained in three different wingspan locations for the mesh. The objective was to confirm that the limiting streamline contours are coherent with the pressure distribution found in the upper and lower surface of the wing. In figure 5.17, the schematic of the three different sections where the pressure distribution and

Figure 5.16: Streamlines of the flow in the upper surface of the wing for AoA: 18^◦.

respective limiting streamline contour were analysed for the fine mesh are shown. It is observed that the section that has higher lift, i.e. higher pressure difference, is the section located at z = 2.5 m, which is in fact the section where the flow is attached up to a more extensive chord. The suction peak differs for the three different sections due to the wing washout, i.e., at the root the angle of attack of the wing is higher and at the tip it is lower. In fact, the suction peak is lower at the section located in the plane z = 1.5 m and higher at the plane z = 4 m which is close to the tip and has a lower AoA.

(a) Schematic of the different sections in the wing for which the pressure distributions were computed.

(b) Negative pressure coefficient as a function of the normalized chord at AoA = 18^◦in the planes z = 1.5, 2.5 and 4 meters

Figure 5.17: Pressure distribution of three different sections on the wing at an AoA = 18^◦and respective schematic.

As complement to streamline contours, contours of the skin friction coefficient were also analysed.

The skin friction coefficient (Cf) is the non-dimensional shear-stress at the wall . For a finite wing the separation occurs no longer at point (as it happen for an aerofoil) but rather in a separation line. In the separation line the value ofCf is zero and when the flow is separated the skin friction coefficient is negative due to the presence of adverse pressure gradients. In order to observe where separation of the boundary layer occurred, contours ofC_{f x}were obtained with the CFD-Post. The contours have the same scale.

In figure 5.18, the skin friction coefficient contours are shown for the range of angles of attack studied. The progression of the boundary layer separation is equivalent to the previous section, i.e., the C_f is zero at the wing root for lowerαand it progresses both spanwise and chordwise asαis increased.

Atα= 16^◦theC_fis zero at approximately50%of the upper surface of the wing. It is concluded, that the aircraft is stalled when 50%of the boundary layer is separated.

Figure 5.18: Skin Friction coefficient in the flow direction.

In order to evaluate the quality of the results obtained from the CFD flow calculations, the be- haviour of wool tufts in the upper surface of the wing were monitored. As mention in section 4.2.4 the left wing of the Slingsby Firefly was covered by wool tufts of 15 cm length and 24 cm apart. The aircraft was stalled and the video imagery of the wool tufts were recorded. The flow begins to separate at the root after the stall warning is activated, however buffet only began after 4 seconds of stall warning. As the airspeed is reduce to the stall speed the AoA increases and the flow separation region follows a similar separation pattern as found in the limiting stramline contours in the CFD flow calculation.The separation progresses both spanwise and chordwise as expected for the same previous reasoning stated.

The behaviour of the wool tufts during the pre-stall and stall condition for the four seconds duration of buffet are shown in Appendix B.1. The images result from post-processing of the video recorded, the last image is just before the wing dropped and the aircraft stall is recovered before entering a spin.

Similar patterns, to the ones obtained from the CFD calculations, of the boundary layer separation were observed in the flight test. The stall cells visualization was not possible however, similar regions of attached and separated flow were observed. Although the same patterns occurred, their relative

localization was also verified. Figures 5.19, 5.20 and 5.21 show the comparison between the wool tufts flight test and the CFD limiting streamlines contours pattern and respective localization on the upper surface of the wing. It was observed that that the patterns of the flow separation and respective locus were similar between the numerical and experimental approach.

Figure 5.19: Flow visualisation of the flow with an AoA of 10^◦and equivalent flight test image.

Figure 5.20: Flow visualisation of the flow with an AoA of 14^◦and equivalent flight test image.

Figure 5.21: Flow visualisation of the flow with an AoA of 18^◦and equivalent flight test image.