2.1 Pressure measurements setup

The wing and boom surfaces were equipped with a total of 256 surface static pressure taps distributed as shown in Fig. 4. The pressure taps on the wing surface were positioned in four belts equipped with 40 taps each. Pressure measurements were performed using eight low-range 32-port ESP-32HD pressure scanners embedded inside the wing model, acquired using DTC-Initium pressure scanning system. Acquisition time was 10 seconds for each test individual test point.

Figure 4. Scheme of pressure measurements setup.

2.2 Infrared Thermography measurements setup

The baseline wing aerofoil section was designed to achieve a certain amount of laminar flow. A main benefit of laminar flow is reduced drag, which translates into reduced power requirements and/or increased range. Due to the complex interaction between the periodic blade wake and the wing aerofoil, with different regions of upwash and downwash, it’s extremely difficult to reliably predict the actual amount of achievable laminar flow. During the present wind tunnel test IRT was used to evaluate quantitatively the amount of laminar flow for the baseline aerofoil section and also for the wing aerofoil under the influence of the propeller at different thrust conditions. The IRT instrumentation consisted of a FLIR A6751 LWIR Camera equipped with a Strained-Layer Superlattice sensor with $640 \times 512{\rm{px}}$ resolution (square pixel size of 15 ${\rm{\mu m}}$ edge). The thermal sensitivity of the sensor is below 45mK. The camera was robustly attached to the wind tunnel floor by means of a metallic support structure (see Fig. 2). A set of markers made of silver conductive paint dots were applied to the upper surface of the wing model in order to have known locations on the curved surface useful to map the images on the actual wing geometry. The markers were smoothed to avoid any roughness that could accelerate or trigger premature transition. Due to its size and the possible interference with other sensors, the model isn’t equipped with an internal heating system, thus the test procedure was to heat up the tunnel free-stream air with the heat exchanger disabled, resulting in the model being cooler than the air. A small temperature difference of 2°C to 3°C between the model and free-stream flow is sufficient to capture the differences in heat exchange between the laminar and turbulent BLs. Two different fields of view over the wing upper surface, above and below the boom, were considered for the IRT measurements. The measurements included conditions with different wing angles of attack, tilter propeller rotational speeds, blade collective angles and nacelle angles, as well as wind tunnel free stream velocities. For each model test condition and field of view a series of $500$ individual images were acquired with an acquisition frequency of 100Hz and an exposure time (integration time) of 131.5 ${\rm{\mu s}}$ .

3.1 Pressure measurements

This section presents a selection of pressure measurements obtained during the test campaign. Pressure measurements included a wide set of test conditions used to evaluate the effect of the tilter propeller interaction with the wing performance in cruise-flight condition. For all test conditions the lifter propeller was kept stowed. Test points included different sweeps of wing angle-of-attack, tunnel speed, propeller tilt angle, blade collective angle and propeller rotational speed. In the following, test results obtained for a representative cruise flight condition, i.e. with zero deflection of tilter propeller shaft angle and ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ , are presented and discussed.

Figure 5 shows the pressure coefficient distributions measured on the four wing belts at a constant propeller rotational speed of 1,000rpm and at wing angles of attack ranging from −2.5 ${{\rm{\;}}^ \circ }$ to 5 ${{\rm{\;}}^ \circ }$ .

Figure 5. Static pressure measurements results over the wing with powered tilter propeller (1,000rpm) at different angles of attack, ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ .

The pressure distribution on the wing changes along the span of the model as a result of the induced effect of the propeller slipstream on the wing as well as the propeller sense of rotation. This is clearly visible from the comparison of the chord-wise pressure coefficient distributions measured over the four instrumented wing sections for a given angle-of-attack. For instance, the propeller induces a local upwash that has the effect of increasing the local angle-of-attack on belts #1 and #2. On the other hand, belts #3 and #4, located on the other side of the boom, are subject to a downwash that decreases the local angle-of-attack. The effect is further magnified by comparing results over the outer belts #1 and #4, where the vertical velocity component induced by the propeller is higher compared to the inboard belts.

Also discernable in the pressure distributions is the effect of angle-of-attack on the aerofoil upper surface favorable pressure gradient. The magnitude and extent of the favorable pressure gradient has a direct effect on the BL transition location. As expected, as the angle-of-attack of the model increases, the favourable pressure gradient near the leading-edge of the wing suction side (i.e. the upper surface of the wing) decreases becoming less favourable. As this favourable gradient decreases, one would expect the BL transition location to move forward towards the leading edge. The BL transition location depends primarily on the characteristics of the wing aerofoil section and it is driven by the angle-of-attack of the wing. Other factors may also influence the transition location including the propeller interaction effects and tunnel flow quality.

The pressure measurements on the four wing belts shown in Fig. 5 have been integrated to obtain the corresponding sectional lift coefficients ${C_l}$ in order to quantify the effect of the tilter propeller slipstream on the wing. The results are scaled with respect to a reference sectional lift coefficient ${\bar C_{l,np}}$ obtained by averaging the local lift coefficients on the four belts measured with the propeller turned off at fixed wing angle-of-attack ( $\alpha = {2.5^ \circ }$ ) and at a given wind tunnel freestream velocity ( ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ ). The ratio between the sectional lift coefficient ${C_l}$ and the reference lift coefficient ${\bar C_{l,np}}$ is shown in Fig. 6 as function of angle-of-attack at fixed rotational speed of the propeller for each belt.

Figure 6. Ratio between sectional lift coefficient ${C_l}$ and the reference sectional lift coefficient ${\bar C_{l,np}}$ over the wing with powered tilter propeller (1,000rpm) as function of angle-of-attack at different spanwise positions, ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ .

From Fig. 6, the upwash induced by the propeller on the left side of wing and experienced by the sections corresponding to belts #1 and #2 has the effect to increase the lift on this portion of the wing with respect to the opposite side where the propeller rotation induces a local downwash (belts #3 and #4). The lift variation between belts #1 and #2 in the upwash region of the wing is higher than the one observed between belts #3 and #4. Furthermore, sectional lift coefficients evaluated on belts #3 and #4 are fairly similar throughout the whole range of angles of attack tested. Since the wing is straight wing, changes in the pressure distribution and local lift coefficient are due to the 2D aerofoil contour and angle-of-attack. Any spanwise differences in the distribution and lift coefficient are attributable to the propeller and its sense of rotation, which induces upwash/downwash velocities on the sections of the wing model depending on their position relative to the propeller axis.

The non-uniform lift distribution promoted over the wing span by the propeller slipstream is clearly highlighted in Fig. 7, where the ratio of sectional lift coefficients ${C_l}/{\bar C_{l,np}}$ is shown as function of the belt spanwise position at a fixed propeller rotational speed and for different angles of attack. Figure 7 shows that the wing sectional lift coefficient monotonically decreases going from belt #1 to belt #4. This behaviour is fairly consistent for all the angles of attack analysed. A net decrease of sectional lift coefficient occurs in the upwash region of the wing between belt #1 and #2 while in the downwash region between belt #3 and #4 the sectional lift coefficient reaches a plateau, particularly at higher angles of attack.

Figure 7. Ratio between sectional lift coefficient ${C_l}$ and the reference sectional lift coefficient ${\bar C_{l,np}}$ over the wing with powered tilter propeller (1,000rpm) as function of spanwise coordinate at different angles of attack, ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ .

Figure 8 shows a comparison between the pressure distributions measured on the four belts of the wing with both the propeller spinning at different rotational speeds and turned off, at fixed wing angle-of-attack ( $\alpha = {2.5^ \circ }$ ) and at a given wind tunnel freestream velocity ( ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ ). The powered test cases are presented with the tilter propeller operating at a rotational speed ranging from 700 to 1,100rpm to better illustrate the propeller rotational speed effects on the pressure distributions.

Figure 8. Static pressure measurements results over the wing with and without powered tilter propeller (700–1,100rpm) at $\alpha = {2.5^ \circ }$ , ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ .

The pressure distributions measured on the wing belts show similar trends at high propeller rotational speed settings (1,000 and 1,100rpm). The effect of the propeller slipstream on the pressure distributions is noticeable on the outer belts (#1 and #4) when the powered cases are compared to the unpowered one. The effect of the propeller slipstream decreases on the inner belts (#2 and #3) where the differences with respect to the unpowered case are fairly small. At low propeller rotational speed setting (700rpm) the effects induced by the propeller slipstream on the wing model changes considerably with respect to the high rotational speed cases. The combination of propeller rotational speed and wind tunnel freestream velocity ( ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ ) produces an inflow that drives the propeller in windmill state. The pressure distributions measured on the wing in this condition show an opposite trend when compared to the high rotational speed cases and are qualitatively similar to the pressure distributions measured with the propeller turned off.

The effects produced by the propeller slipstream on the wing are clearly illustrated in Fig. 9 where the ratio of sectional lift coefficients ${C_l}/{\bar C_{l,np}}$ on each belt is shown as function of their spanwise coordinate with both the propeller spinning and turned off. Sectional lift coefficients ${C_l}/{\bar C_{l,np}}$ are compared at a fixed angle-of-attack ( $\alpha = {2.5^ \circ }$ ) and considering the same range of propeller velocities between 700 and 1,100rpm. Figure 9 confirms that the flow around the wing is highly influenced by the propeller slipstream, which provides a local opposite sectional lift variation with respect to its rotational axis. On the other hand, when the propeller is turned off, the lift distribution across the wing span shows a fairly flat behaviour, as expected. This proves that propeller rotational speed and its sense of rotation are directly responsible for the asymmetry in the local flow around the portion of the wing blown by the propeller wake. The absolute value of the sectional lift coefficient variation provided at propeller rotational speeds equal to 1,000 and 1,100rpm is quite higher in the upwash region (belts #1 and #2), as previously observed. The effect on local wing airloads is more apparent at 1,100rpm. At 700rpm the propeller is in windmill state and an opposite variation of local sectional lift with respect to propeller rotation axis can be observed.

Figure 9. Ratio between sectional lift coefficient ${C_l}$ and the reference sectional lift coefficient ${\bar C_{l,np}}$ over the wing as function of spanwise coordinate with and without powered tilter propeller at $\alpha = {2.5^ \circ }$ , ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ .

3.2 Infrared Thermography measurements

The images displayed in Fig. 10 illustrate the surface temperature distribution on the suction side of the wing as a function of angle-of-attack. The images were obtained from the IRT survey carried out on the wing model with the tilter propeller removed from the boom and at fixed wind tunnel freestream velocity ( ${U_\infty } = 40\frac{{\rm{m}}}{{\rm{s}}}$ ). The figure shows the average intensity distributions over the wing model at different angles of attack $\alpha $ ranging from $\alpha = - {2.5^ \circ }$ (leftmost image) to $\alpha = {10^ \circ }$ (rightmost image). The $x$ - and $y$ -axes represent the chord and span of the wing, respectively, whereas the surface temperature is given as a grey scale distribution averaged over the $500$ individual recordings. The images were mapped onto a 3D grid of the wing’s surface using the in house DLR software ToPas [Reference Klein, Engler, Henne and Sachs18].

Figure 10. Average intensity distributions at different pitch angles $\alpha $ and an inflow velocity of ${U_\infty } = 40$ m/s without the propeller mounted. Minor disturbances in the shape of the transition line due to the presence of surface static pressure taps (“wedges”) are highlighted in red.

Figure 10 shows that the temperature at the leading-edge of the wing is less than that at the trailing-edge (darker gray scale level). This temperature difference is the effect of the flow behaviour on the wing’s surface. For a laminar region, turbulent mixing is reduced and thus the heating of the wing’s surface is reduced, while in a turbulent region the heat exchange is increased and thus the wing is heated more effectively. This effect can be exploited to measure the position of BL transition with IRT. Within the global measured field-of-view (FoV) depicted in Fig. 10 the temperature distribution is symmetric with respect to the $y$ -axis, which is expected based on the straight wing symmetric geometry (above and below the boom). The images displayed in Fig. 10 show that the transition location moves towards the leading edge as the angle-of-attack of the wing increases, as expected. The influence of the boom can be seen in the region around the $y$ -axis ( $y/c = - 0.1 \ldots 0.1$ ), but since the map temperature is almost constant, no valuable information can be inferred in this portion of the wing. Instead the influence of pressure taps on the wing’s surface as spanwise “wedges” can be observed as narrow areas of local forward movement of the transition position (see the case $\alpha = {7.5^ \circ }$ at spanwise positions $y/c = \pm 0.55$ ). The two dimensionality of the transition front and its movement with angle-of-attack indicates good tunnel flow quality and the absence of wall effects on the spanwise behaviour of the front.

Figure 11(a) illustrates the intensity distributions for all measured angles of attack along the wing chord at a fixed spanwise position ( $y/c = 0.4$ ) and corresponding to the orange lines shown in Fig. 10. The curves are at slightly different intensity levels due to differences in the overall temperature change of the wing model during different runs. Nevertheless, a noticeable increase in the measured intensity along the chord of approximately $20$ counts is evident for all angles of attack considered. This increase denotes the location of BL transition. According to Weiss et al. [Reference Weiss, Gardner, Klein and Raffel19], the position of 50% turbulence intermittency is found at the highest gradient of the temperature in chordwise direction. The corresponding spatial gradient curves for the current test cases are reported in Fig. 11(b), where a distinct maximum can be observed on each curve. The respective positions are marked as colored dots in both Fig. 11(a) and (b). These results can be used as a benchmark for the interpretation of the transition under the influence of the propeller.

Figure 11. Left: Chord wise intensity profiles extracted from the data of Fig. 10 at $y/c = - 0.4$ . Right: First derivative of the intensity profiles shown on the left.

Figure 12. Transition locations as function of angle-of-attack for different inflow velocities measured with IRT and predicted with XFOIL.

The detected transition locations are plotted against the wing angle-of-attack for both wind tunnel freestream velocities ( ${U_\infty } = 40\frac{{\rm{m}}}{{\rm{s}}}$ and ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ ) in Fig. 12 as dots. The black dots are in the same locations as depicted in Fig. 11. In addition to the IRT results, predicted transition locations from 2D XFOIL calculations with an ${N_{{\rm{crit}}}} = 9$ are shown as solid lines. The results confirm the expected trend of the BL transition location that moves towards the leading edge as the angle-of-attack increases. The effects of the wind tunnel freestream velocity (increased Reynolds number) are also consistent with the data collected. The previous 2D XFOIL analysis was also repeated at a freestream velocity of ${U_\infty } = 100\frac{{\rm{m}}}{{\rm{s}}}$ , significantly higher than the maximum freestream velocity reached during the test ( ${U_\infty } = 50\frac{{\rm{m}}}{{\rm{s}}}$ ). Such analysis was carried out to further assess the effects on the BL transition location of an increased inflow velocity due to the presence of the propeller in front of the wing model. As it will be shown later in this paper, the comparison between the measured BL transition location under the influence of the propeller and the 2D XFOIL results points to a possible combination of local induced inflow angle and propeller induced turbulence.

The intensity distributions in Fig. 13 are taken at an angle-of-attack $\alpha = - {2.5^ \circ }$ under the influence of the propeller. The image in Fig. 13(a) shows the previously discussed case without the propeller for comparison. The images in Fig. 13(b)–(e) result from different rotational speeds of the propeller. From Fig. 13(b)–(d) it can be seen that the propeller triggers an earlier onset of transition. Furthermore, the impact of the propeller blade tip vortices is clearly visible at $y/c$ = $ \pm $ 0.7 for 1,000 and 1,200rpm as transition cones starting at the leading-edge. The earlier onset of transition is a result of the increased turbulence in the propeller wake and the increased Reynolds number compared to the undisturbed wind tunnel flow. The propeller results in Fig. 13(b)–(d) show a rather symmetric transition region for positive and negative $y/c$ . This points to a comparably low influence of the direction of rotation of the propeller (upwash/downwash). Furthermore, it can be seen that the transition line for the propeller cases is not as distinct as in the no propeller case.

Figure 13. Average intensity distributions without the propeller (a) and with the propeller spinning at different rotational speeds (b)–(d), all at $ = - {2.5^ \circ }$ and $U\infty = 40$ m/s. Footprints of the propeller blades’ trailing vortices are highlighted in red.

To further investigate the transition topology under the influence of the propeller, intensity profiles similar to those in Fig. 11 at $\alpha = - {2.5^ \circ }$ , but related to different rotational speeds are shown in Fig. 14 at a spanwise location at y/c=-0.4. From the resulting intensity levels it can be seen that the temperature increase in chordwise direction is not as distinct as in the case without the propeller. Therefore, the maxima of the gradients are less prominent and, thus, the measurement of the transition location is more ambiguous.

Figure 14. Left: Chordwise intensity profiles extracted from the data of Fig. 13 at $y/c = - 0.4$ . Right: First derivative of the intensity profiles shown on the left.

Figure 15. Spanwise distribution of detected transition locations with propeller spinning and turned off at different angles of attack and rotational frequencies at a wind tunnel freestream velocity of ${U_\infty } = 40$ m/s.

The process demonstrated in Fig. 14 can be repeated for every row in the $y/c$ -direction to obtain the spanwise distribution of the transition location. Figure 15 shows the comparison of these spanwise distributions for different angles of attack (a-d) and propeller rotational speeds at a fixed wind tunnel freestream velocity of ${U_\infty } = 40\frac{{\rm{m}}}{{\rm{s}}}$ . The transition location distributions obtained with the propeller spinning are compared with the the unpowered case in each figure. The spanwise distribution of the transition location for the test cases shown in Fig. 13 at $\alpha = - {2.5^ \circ }$ are presented in Fig. 15(a). Detected locations in the region of overlap with the boom ( $\left| {y/c} \right| \lt 0.19$ ) were masked out. The comparison between the powered and unpowered cases clearly shows that the BL transition is not affected by the presence of the propeller in the portions of the wing that fall outside the propeller slipstream ( $\left| {y/c} \right| \gt 0.85$ , e.g. not directly hit by the propeller wake). A different behaviour is observed in the wing portion that is affected by the propeller wake. As expected, in this region the BL tends to transition earlier than when the propeller is turned off. Due to the less prominent step in the chordwise intensity distribution (see Fig. 14) a scatter in the detected transition location occurs and the detected transition location is not as defined as outside of the propeller wake. As documented by Miley [Reference Miley, Howard and Holmes20] and Howard [Reference Howard21], both flight test measurements and wind tunnel experiments carried out to investigate the effects of the propeller slipstream on a wing have shown that the BL cycles between a laminar and a turbulent state at the propeller blade passage rate. As the BL on a given point on the wing contour goes through distinct phases of turbulent, reverse-transitional, and laminar behaviour, the surface temperature of the wing model is also affected. This in turn leads to smoother temperature distributions as the images recorded by the IRT camera are temporally smoothed by the heat capacity of the wing surface and the integration time of the camera exposure. Different phenomena may cause the earlier onset in the portion of the wing blown by the propeller wake. Both an increased inflow velocity and turbulence level in the propeller wake may lead to an earlier onset of the BL transition on the wing compared to the unpowered cases. Based on the 2D XFOIL calculations at ${U_\infty } = 100\frac{{\rm{m}}}{{\rm{s}}}$ the influence of an increased inflow velocity (and thus increased Reynolds number) alone can be estimated. The XFOIL calculated transition position occurs at higher $x/c$ (further downstream) than documented in the IRT measurement in the region of the propeller wake. This implies that the earlier transition onset in the region of the propeller wake cannot be ascribed to the increased inflow velocity alone. At the slipstream boundaries ( $\left| {y/c} \right| \approx 0.7$ ) the aforementioned cones triggered by the propeller blade tip vortices are observed.