Slatless configuration

Before performing the wind tunnel tests, numerical simulations have been performed to help designing the flow control devices. Starting with the slatless configuration, the question is: at which location to put the blowing slot to maximise CLmax? So, first, a parametric study of the ef-fect of the slot location on lift has been performed. In order to avoid remeshing the whole configu-ration for each slot location, the overset grid method has been used (see Fig. V.18 (left)). Fig.

V.18 (right) shows the effect of the slot location on lift. The first lesson learnt is that the region close the leading edge, where there is the Cp peak (see Fig. V.19 (left)), must be avoided because the local velocity is very high (≈ 300 m.s^-1) so it requires very high blowing velocity to have a velocity ratio over one. The second lesson learnt is that the slot must be located before the separa-tion point or possibly a little bit downstream. The optimum slot locasepara-tion is here at x/c = 4%.

Fig. V.18: Overset grid strategy used to add the slot in the background mesh (left) and slot loca-tion effect on the lift coefficient (right) (α = 15°, Cµ = 2.5% and blowing angle 20° wrt to local

tangent).

The effect of the continuous blowing slot on the wall pressure distribution is plotted in Fig. V.19 (right). Compared to the uncontrolled (left figure), the local velocity at the leading edge, which was already very high, has been increased to Mis = 1.36 (Cp = -23.5). This corresponds to the so-called 0.7 vacuum limit before stall according to Mayer [40]. He showed in 1948 in his experiments that there is no attached flow when Cp.M² < -1. This has also been verified by Wild [56]. It corresponds to Cpmin = -25 for M = 0.2. This limits the performance of active flow control because this Cpmin value limits the maximum lift coefficient achievable.

Fig. V.19: Wall pressure distribution on the slat-less configuration without control (left) and with a continuous blowing slot at x = 0.02 (right).

Fig. V.20 shows the effect of the continuous blowing slot on the lift curve. Despite an al-ready very large value of the Cµ coefficient (2.5%), the stall has only been delayed by 6° and the control is not able to compensate completely the slat suppression since the stall angle of the base-line configuration is at 25°.

Fig. V.20: Lift curve for the three-element configuration (in red), the slat-less one without control (in blue) and with control by a continuous blowing slot (in cyan).

To try to understand the limitations of active flow control at the leading edge, the lift is plotted as function of the freestream Mach number for different angles of attack on a case with a continuous blowing slot at the leading edge (see Fig. V.21). One can remark that the higher the freestream Mach number, the earlier the stall. This can be explained by the appearance of a shock upstream of the slot (see the Mach number field in Fig. V.23). As previously said, the local Mach number is equal to 1.36 and it is known that for this value the shock/boundary layer interaction leads to a separation at the shock foot which is visible in Fig. V.23. Fig. V.21 shows that decreas-ing the freestream Mach number allows delaydecreas-ing more stall and consequently to increase more CLmax. For example, at M∞ = 0.115, CLmax is equal to 4.6 and stall occurs at 24° like the three-element configuration which was the objective of the control. Unfortunately, civil aircraft have a landing Mach number around 0.2 and not 0.115. So, in summary, compressibility effects limit the performance of active flow control at the leading edge.

Fig. V.21: Lift coefficient as function of the freestream Mach number for several angles of attack (with control by a continuous blowing slot with qm= 100 g.s^-1 per meter span).

Fig. V.22 shows lift curves without control and with control by continuous blowing. In the baseline case, there is an abrupt stall whereas with control, the stall is smoother. This means that the leading edge stall has been replaced by a trailing edge stall. The leading edge blowing stall is not able to prevent the airfoil from a trailing edge separation (visible in the velocity fields). One should add a second blowing slot, at mid-chord for example, to suppress this trailing edge separa-tion. But, in practice, this is not possible because the fuel tank is located between 25% and 75% of the chord.

Close to the stall angle, the stagnation point is located on the lower side of the airfoil.

Since the velocity ratio is important as shown for example in the definition of the CBLC coefficient, the idea was to put a blowing slot on the lower side between the stagnation point and the Cp peak.

Moreover, as explained in the introduction, active flow control at the leading edge could be used to replace slats which are incompatible with laminarity in cruise. So, by putting the blowing slot on the lower side, it will not trigger the laminar/turbulent transition on the upper side in cruise since the stagnation point will move upstream of the slot. Fig. V.22 shows the effect on the lift curve. The lower side slot (in brown) has nearly the same lift curve as the upper side slot with the same mass flow rate (in red). So, putting the slot on the lower side instead of the upper one does not decrease the performance. It does not either improve CLmax even if the velocity ratio is higher.

This is also due to the shock/boundary layer interaction which limits the performance. Neverthe-less, as explained in the synthetic jet section (1.1), the actuator performance is limited by the suc-tion phase. By putting, the synthetic jet on the lower side where the local velocity is lower and hence the static pressure is higher, the cavity filling will be improved and consequently the blow-ing phase and the actuator momentum flux.

Fig. V.22: Lift curves with a continuous blowing slot for different slot widths L and mass flow rates Q and slot location (upper or lower side) (M∞ = 0.2).

Fig. V.23 shows the Mach number field for a tangential slot with a blowing angle of 30°.

For slot widths larger than 0.1% of chord, a very small separation bubble (which would not be visible in the experiment) is present downstream of the slot. This separation induces a low veloci-ty deficit in its wake. The peak velociveloci-ty is not located close enough to the wall so it is not as effec-tive as it could be. So, by modifying the slot geometry in order to be more tangential, the separa-tion bubble has been suppressed. The effect on the lift curve is visible in Fig. V.24 (top) by com-paring the black curve with the purple one. The lift gain is nearly doubled for large Cµ values which is not negligible. Fig. V.24 (top) also shows that the 0.25 mm side slot is able to provide the same lift gains as the 0.5 mm wide one but for a mass flow rate divided by two. Nevertheless, the 0.25 mm wide slot is limited to a maximum Cµ value of 2% and consequently a lift gain of 10%. In order to reach the objective of 25% without increasing the slot width, a slot with a nozzle is used with jet Mach numbers of 1.4 (= √2) and 2. This nozzle allows reaching the objective of a CLmax increase of 25%. Then, since the 0.25 and 0.5 mm curves do not collapse, the lift gain is plotted as function of the velocity ratio VR = Uj/Uloc (where Uloc is the local velocity) instead of Cµ. This allows collapsing all curves. This means that, on the present test case, the slot width seems to have no effect on the lift gain. This is quite strange since it means that a wide slot would have the same effect as a very thin slot or maybe there is plateau. But, this would require further studies to be confirmed on other cases. A slot width of 0.25 mm corresponds to a non-dimensionalised with d/δ = 0.53.

Fig. V.23: Mach number field in the leading edge area.

Fig. V.24: Lift gain as function of the momentum coefficient Cμ (top) and velocity ratio VR (bot-tom).

Fig. V.25 (left) shows the effect of Cµ on the lift gain for two different leading edges: one laminar (in red) and one turbulent (in blue). The slot is located at the same location with the same width. The lift gain is lower for the laminar leading edge than for the turbulent one. This is due to the fact that a laminar leading edge has a smaller radius than a turbulent leading edge and conse-quently the local velocity is higher (270 m.s^-1 instead of 200 m.s^-1). This limits the control effi-ciency and for a given Cµ value, the lift gain is smaller. To take into account this higher local

ve-locity, the Cµ definition is modified by replacing the freestream velocity by the local one at the slot location (see Fig. V.25 (right)). This modification enables to collapse the two curves.

Fig. V.25: Lift gain as function of the momentum coefficient and a modified momentum coeffi-cient for a turbulent airfoil (in blue) and laminar airfoil (in red).

To improve active flow control effect, another idea is to decrease the local velocity, for example by using a droop nose (configuration n°3 in Fig. V.15).