Study of a Swept Wing with Leading-Edge Ice Using a Wake Survey Technique

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1 51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition 7-1 January 213, Grapevine (Dallas/Ft. Worth Region), Texas AIAA Study of a Swept Wing with Leading-Edge Ice Using a Wake Survey Technique Jeff M. Diebold * and Michael B. Bragg University of Illinois at Urbana-Champaign, Urbana, IL 6181 Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ The aerodynamics of swept wings with leading-edge ice is very complex and not fully understood. Previous swept wing icing studies have only provided force balance measurements and/or qualitative flow visualization. In order to more fully understand the complex aerodynamics it is necessary to understand how the ice influences the drag breakdown and the spanwise distributions of lift and drag. This paper utilizes 5-hole probe wake surveys and a far-field analysis to decompose the drag into profile and induced drag components and to determine the spanwise distributions of lift and drag. It is shown that the leading-edge ice primarily influences the profile drag and the induced drag is relatively unaltered. Features observed in the lift and drag distributions are related to features seen in the wake as well as in surface oil flow visualization. Nomenclature c(y) = Chord distribution C d = Sectional total drag coefficient C di = Sectional induced drag coefficient C dp = Sectional profile drag coefficient C l = Sectional lift coefficient C D = Total Drag coefficient C Di = Induced drag coefficient C L = Lift coefficient D = Total Drag D i = Induced drag D p = Profile drag L = Lift P = Static pressure P t = Total pressure in the wake P t = Freestream total pressure u = Streamwise velocity in the wake u b = Wake blockage velocity u = Perturbation velocity u* = Artificial velocity U = Freestream velocity v = Spanwise velocity in the wake w = Normal velocity in the wake x = Streamwise direction y = Spanwise direction (positive from root to tip) z = Normal direction (positive from lower to upper surface) α = Model angle of attack (measured at the root) ψ = Stream function in the transverse plane ξ = Streamwise vorticity * Graduate Research Assistant, Department of Aeronautical and Astronautical Engineering, Member AIAA Professor of Aerospace Engineering, Interim Dean of College of Engineering, Fellow AIAA 1 Copyright 213 by J. Diebold and M. Bragg. Published by the, Inc., with permission.

3 2. Wake Survey Analysis By analyzing the control volume shown in Fig. 1 the following expressions for lift and drag can be derived, as shown in Brune. 11 Note that lift is in the positive z-direction. (1) (( )) ( ) ( ) (2) Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ While these two expressions are exact, they are not practical from an experimental point of view. Calculating the lift requires integrating the pressure over the sidewalls, and both expressions contain integrals that must be performed over the entire downstream plane S 2. A practical set of equations will consist only of wake integrals such as the first integral in Eq. 2, where outside of the wake and the integrand is zero. A detailed derivation of the wake survey equations used in this paper will not be given but can be found in numerous sources. 1,11,12 The derivation follows the work of Betz 13 and Maskell 14 and the final expressions are given here. The final expressions for lift and drag require measurements of the pressure and velocity components in the viscous wake. One of the main advantages of the wake survey as opposed to force balance measurements is the ability to decompose drag into profile and induced drag components. The wake integral expressions for the two components of drag are given by Eqs. 3 and 4. Fig. 1 Control volume and coordinate system used in wake survey. (( ) ( )( ( ))) (3) (4) The total drag was the sum of the profile and induced drag components. In Eq. 3, and are known as the artificial velocity and wake blockage velocity, respectively. The artificial velocity was introduced by Betz 13 and is defined in Eq. 5. Notice the artificial velocity is equal to the freestream velocity outside of the viscous wake. ( ) (5) The wake blockage velocity was interpreted by Maskell 14 as a small correction due to the effect of tunnel walls. ( ) (6) The profile drag only requires measurements of the total pressure and axial velocity within the wake, outside of the wake the integrand is zero. In Eq. 4, for the induced drag, and are the streamwise vorticity and the transverse stream function, respectively. The vorticity is defined as: (7) 3

4 The transverse stream function satisfies the following Poisson equation. Requiring the tunnel wall to be a streamline in the transverse plane provides the following boundary conditions for on the tunnel walls. ( ) (9) Equation 4 for the induced drag is a wake integral, because the vorticity ( ) is zero outside of the wake. The lift is given by the following expression, where y is the distance measured from the tunnel floor. (1) (8) Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ The sectional lift coefficient can be determined by assuming a planar wake and applying the Kutta-Joukowski theorem. 12 ( ) ( ) ( ) (11) Where c(y) is the local chord measured in the streamwise direction. The integrands of Eqs. 1 and 11 are zero outside of the wake. 3. Experimental Methods This section will discuss the relevant experimental techniques used in this study. The model and ice shape simulation are described in Diebold et. al. 9 and will therefore only be briefly discussed here. 3.1 Wind Tunnel Wind tunnel tests were performed in the University of Illinois subsonic wind tunnel. The wind tunnel is of the open-return type with a rectangular test section measuring 2.8-ft by 4-ft. The maximum speed in the empty test section was approximately 165 mph (242 ft/s), this corresponds to a Reynolds number of 1.5x1 6 /ft. The turbulence intensity in the test section was approximately.1% over the entire operating range. 3.2 Model For this study a modified version of the Common Research Model (CRM) was used. The CRM was designed by Vassberg et. al. 15 for the 4 th AIAA Drag Prediction Workshop. It was chosen for the original swept wing icing study by Diebold et. al. 9 because in addition to representing the wing of a modern commercial airliner, the full geometry as well as experimental and computational data are publicly available. The model used at the University of Illinois was of the reflection plane type and is shown in Fig. 2. The geometry was modified slightly from the original CRM to ease construction; these modifications are discussed in Diebold et. al. 9 The relevant geometric Fig. 2 Wind tunnel model. parameters are given Table 1. Table 1 Geometric features of the swept wing model. AR LE Sweep MAC (ft) Semispan (ft) Taper Ratio

7 Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ α = 2.3º α = 4.4º α = 6.5º α = 8.6º α = 1.7º Fig. 4 Surface oil flow for the clean wing at several angles of attack. Re = 6x1 5 α = 2.3º α = 3.3º α = 4.4º α = 5.5º α = 6.5º Fig. 5 Surface oil flow for the iced wing at several angles of attack, with the reattachment line highlighted. Re = 6x1 5 7

8 Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ C L Figure 6 compares the total lift and drag measured by the force balance and wake survey for the clean and iced wing at Re = 6x1 5. Comparing the force balance data of the clean and iced wing it can be seen that below stall the ice caused a small decrease in lift for a given angle of attack and a significant increase in drag. The increase in the minimum drag coefficient due to the ice was 81.1% while at a lift coefficient of.5 the increase in the drag coefficient was 69.7%. The clean wing stalled at an angle of attack of approximately 9.7º while the iced wing stalled at approximately 6.º. The ice significantly changed the behavior of the lift curve after stall. For the clean wing the lift was nearly constant after stall, but the lift of the iced wing continued to increase but the slope was reduced. Comparing the force balance measurements and the wake survey results it can be seen that the wake survey method was able to accurately determine the lift and drag. The average error, relative to the balance, in the lift calculated from the wake data was 3.5% and 2.3% for the clean and iced wing, respectively. For the drag, the average error was 14.6% and 13.3% for the clean and iced wing, respectively. While the error in the drag was relatively large, it can be seen in Fig. 6b that the drag curve determined from the wake survey follow the curve from the force balance measurements very well, but the wake survey consistently under predicted the drag. A possible explanation for this was that the inboard 1% of the wake was not surveyed and therefore the drag contribution from this region would not have been captured. Overall the wake survey was able to accurately capture both the lift and drag including the effects of stall Fig. 6 Comparison of lift (a) and drag (b) from the balance and wake. Re = 6x1 5 C D Figure 7 shows the lift coefficient versus the decomposed profile and induced drag components for the clean and iced wing. The lift coefficient and drag coefficients were determined from wake survey data. The results show that, for a given C L, the profile drag of the iced wing was significantly higher than for the clean wing. Over the range of C L s shown the ice shape more than doubled the profile drag. This was expected given the large region of separated flow behind the ice shape seen in the oil flow visualization. It can also be seen in Fig. 7 that as C L was increased, the profile drag of the iced wing increased much faster than for the clean wing. This rapid rise in profile drag was due to the increasing size of the leading-edge vortex on the iced wing as seen in Fig. 5. The induced drag was not significantly influenced by the presence of the ice shape. Figure 7 shows that the curves of C L vs. C Di of the clean and iced wing almost exactly coincided for C L <.55. As a result, for a given C L below iced wing stall, the increase in drag due to the ice shape was nearly entirely due to a rise in the profile drag. Figure 7 also shows that the majority of the drag rise at stall, for the clean and iced wing, was due to a rise in the profile drag. These results show that the primary effect of the ice on the drag was to increase the axial momentum loss as opposed to changing the amount of streamwise vorticity shed into the wake

9 C L Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ C Dp, C Di Fig. 7 Components of drag for clean and iced wing. Re = 6x1 5. Figure 8 shows the spanwise distributions of the lift coefficient as well as the total drag, profile drag and induced drag coefficients for the clean and iced wing at α = 4.4º. Note that because wake surveys were only performed at a few angles of attack, it was not possible to compare lift and drag distributions of the clean and iced wings for the same C L. At α = 4.4º, total C L and C D of the clean wing measured by the force balance were.522 and.27 respectively. For the iced wing at α = 4.4º, total C L and C D were.57 and.42 respectively. Over nearly the entire span, the local lift coefficient was decreased, Fig. 8a, and the drag coefficient was increased, Fig. 8b, by the presence of the ice shape. It can also be seen that the largest differences in the local lift and drag coefficients occurred outboard of approximately midspan. This was a consistent trend seen for all angles of attack; when the ice was present the local aerodynamic penalties were greater on the outboard sections of the wing. As discussed above this was also seen in the oil flow images, Fig. 5, where the size of the recirculation region occupied a larger percentage of the chord on the outboard sections of the wing. It was seen in Fig. 3 that the size of the ice shape relative to the local chord was much larger for the outboard sections of the wing, and this likely plays in important role in the ice shape having a larger effect near the tip. In addition to the relative size of the ice shape the spanwise pressure gradient may be another factor that led to increased performance degradation for the outboard sections of the wing. The spanwise flow, which results from the spanwise pressure gradient, acts as a form of boundary layer suction for the inboard sections making these sections more resistant to separation and possibly promoting reattachment of the separated shear layer. 18 Figures 8c and 8d show spanwise distributions of the profile and induced drag. Consistent with Fig. 7, all of the increase in drag due to the ice shape was a result of an increase in the profile drag. Figure 8d shows that differences in the induced drag distributions of the clean and iced wing were negligible along most of the span except near the tip. Near the tip of the clean wing the change in the lift with respect to the spanwise location was much larger than for the iced wing indicating the clean wing was shedding more vorticity into the wake. The 5HP wake surveys of Diebold et. al. 9 showed that the tip vortex of the clean wing was much stronger than the tip vortex of the iced wing for nearly identical lift coefficients. This resulted in a large peak in the local induced drag of the clean wing near the tip. In addition, the large axial velocity deficit within the tip vortex resulted in peaks in the local profile drag coefficient near the tip for both the clean and iced wings. The axial velocity deficit within the iced wing vortex was larger due to the entrainment of separated flow into the vortex. This resulted in a larger profile drag coefficient near the tip of the iced wing. 9

10 .8 Clean Ice.45.4 Clean Ice C l.4 C d Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ C dp a) Lift Coefficient Clean Ice c) Profile Drag Coefficient C di b) Drag Coefficient Clean Ice d) Induced Drag Coefficient Fig. 8 Spanwise distributions of a) Lift coefficient, b) Drag coefficient, c) Profile drag coefficient and d) Induced drag coefficient for the clean and iced wing at α = 4.4º. Re = 6x1 5. The distributions of total, profile and induced drag, Figs. 8b, c and d, show several local peaks along the span. These peaks were especially prominent for the iced wing. Figure 9 shows how the peaks in the profile drag coefficient of the iced wing at α = 4.4º corresponded to regions of concentrated axial velocity deficit in the wake. The contours in the wake represent normalized axial velocity while the vectors represent the crossflow velocity. Similarly, peaks in the induced drag coefficient corresponded with regions of concentrated axial vorticity as shown in Fig. 1. The peaks in profile and induced drag occurred at the same location. In addition, examining Fig. 8a shows that at these same spanwise locations there is a decrease in the local lift coefficient. This indicates a change in the local circulation occurred explaining the concentrated regions of shed vorticity. It is not surprising that peaks in the drag distributions corresponded to features in the wake since the distributions were determined from wake data, but these figures indicate that something is occurring on the surface of the wing that sheds vorticity and generates a loss of axial momentum. 1

11 Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ Fig. 9 Correlation between regions of concentrated axial velocity deficit and local peaks in profile drag for the iced wing. α = 4º, Re = 6x1 5 Fig. 1 Correlation between regions of concentrated axial vorticity and local peaks in induced drag for the iced wing. α = 4º, Re = 6x1 5 11

12 Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ Figure 11 is an attempt to relate features in the wake, which correspond to peaks in the local drag coefficients, of the iced wing at α = 4.4º to features seen in the oil flow. It can be seen that several of the concentrations of low axial velocity in the wake corresponded with local maxima in the size of the leading-edge vortex. This was observed for all angles of attack for the iced wing, see Diebold. 17 While it is unknown exactly why the size of the leadingedge vortex changed at these locations, it was clear that this had an effect on the local aerodynamic performance. When a local maxima in the leading-edge vortex was encountered, the local aerodynamics of this section were affected more than the neighboring sections because of the larger percentage of the chord being affected by the separated flow. This resulted in a local rise in the pressure drag which is seen as the concentration in low axial velocity in the wake and a peak in profile drag. In addition, because of the larger separated region, the lift at these sections decreased slightly resulting in vorticity being shed into the wake and a peak in induced drag. The low axial velocity region near =.6, and the corresponding local maxima in drag, was observed in all of the wakes for both the clean and iced wing below stall but did not appear to correlate with any feature seen in the oil flow. It is possible that this feature may have originated from the lower surface. Fig. 11 Correlation between features in the wake and the surface oil flow for the iced wing. α = 4º, Re = 6x1 5 The wake survey results were also used to investigate the effects of stall on the clean and iced wing. Figures 12a and 12c compare the spanwise distributions of the clean wing lift and drag for an angle of attack just prior to stall (α = 8.6º) and just after stall (α = 1.7º), Figs. 12b and 12d compare the distributions of lift and drag for the iced wing just prior to stall (α = 5.5º) and just after stall (α = 6.5º). Note the difference in scales for the clean wing figures and iced wing figures. Since the stalling angles of attack are sufficiently different between the clean and iced cases a direct comparison is not as useful. 12

13 Recall from the oil flow images, Figs. 4 and 5, that stall of the clean and iced wing occurred when flow on the outboard stations failed to reattach. The separated flow rolled up to form a leading-edge vortex which was shed into the wake at approximately =.6. The flow inboard of the vortex remained attached while outboard the flow on the surface was fully reversed. The stalled wing flowfields are described in more detail by Diebold et. al. 9 The effects on the lift distribution can be seen in Figs. 12a and b. For both the clean and iced wing the lift on the inboard sections continued to increase through stall, and the lift on the outboard sections decreased significantly due to the separated flow. These results explain the behavior of the lift curve after stall as seen in the force balance measurements in Fig. 6. For both the clean and iced wing the lift did not decrease as a result of stall. For the clean wing, the lift was nearly constant as the angle of attack continued to increase, while for the iced wing the lift continued to increase but the slope of the curve had decreased. The wake survey results show that the inboard sections of the wing are able to compensate for the loss of lift produced by the outboard sections, and prevent the lift from decreasing Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ C d C l Clean, = 8.6 Clean, = a) Clean Lift Coefficient Clean, = 8.6 Clean, = 1.7 C l C d Ice, = 5.5 Ice, = b) Ice Lift Coefficient Ice, = 5.5 Ice, = c) Clean Drag Coefficient d) Ice Drag Coefficient Fig. 12 Comparison of pre and post stall spanwise distributions of a) Clean wing lift coefficient, b) Ice wing lift coefficient, c) Clean wing drag coefficient and d) Ice wing drag coefficient. Re = 6x

14 While not shown here, the force balance measurements also showed an increase in the pitching moment resulting from stall, see Diebold et. al. 9 The mean aerodynamic chord of the model was located at roughly =.4. Therefore, due to the wing sweep, the sections of the wing that stalled were behind the quarter chord of the mean aerodynamic chord as seen in Figs. 12a and b. This resulted in an increased, nose up, pitching moment for the model as a result of stall. The effects of stall on the drag coefficient can be seen in Figs. 12c and d. Inboard of the separated region the distribution of the drag coefficient was relatively unchanged for both the clean and iced wing; however, there was a significant increase in the drag of the outboard sections where the flow had separated. Figures 13a and c show the spanwise distribution of profile and induced drag on the clean wing pre and post stall, and Figs. 13b and d show the corresponding distributions for the iced wing. As was seen in Fig. 7, the drag rise due to stall was primarily due to an increase in the profile drag for both the clean and iced wings. The surface flowfield and wake flowfield of the stalled wings were discussed in detail by Diebold et. al. 9 where it was shown that the separated outboard sections resulted in a large momentum deficit region in the wake. Downloaded by UNIVERSITY OF ILLINOIS on August 14, DOI: 1514/ C di C dp Clean, = 8.6 Clean, = a) Clean Profile Drag Coefficient Clean, = 8.6 Clean, = 1.7 C di C dp Ice, = 5.5 Ice, = b) Ice Profile Drag Coefficient Ice, = 5.5 Ice, = c) Clean Induced Drag Coefficient d) Ice Induced Drag Coefficient Fig. 13 Comparison of pre and post stall spanwise distributions of a) Clean wing profile drag coefficient, b) Ice wing profile drag coefficient, c) Clean wing induced drag coefficient and d) Ice wing induced drag coefficient. Re = 6x

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