AERODYNAMIC PERFORMANCE AND SHEDDING CHARACTERISTICS ON A SWEPT-BACK WING

Transcription

1 162 Journal of Marine Science and Technology, Vol. 19, No. 2, pp (211) AEROYNAMIC PERFORMANCE AN SHEING CHARACTERISTICS ON A SWEPTBACK WING ShunChang Yen* Key words: sweptback wing, vortex shedding, aerodynamic coefficients. ABSTRACT A NACA 12 finite sweptback wing with a sweepback angle of 15 was utilized to investigate the effects of angle of attack (α) and the chord Reynolds number (Re c ) on the vortex shedding and aerodynamic coefficients. A hotwire anemometer was applied to measure the vortexshedding frequency. The projected Strouhal number (St d ) at various angles of attack was determined and discussed. The relationship between St d and α is regressed as: St d =.8 α +.29, for 22 < α < 9. Four characteristic surfaceflow patterns:, leadingedge,, and turbulent were classified by changing α and Re. The behavior of surfaceflow structures significantly affects the lift, drag, and moment coefficients. The lift coefficient (C ) increases with α in the and leadingedge regimes. The maximum increase rate of C with respect to α (d(c )/dα) is 1.52 π/rad. Occurring in the leadingedge regime. However, the maximum increase rate of drag coefficient (C ) with respect to α (d(c )/dα) is.49 π/rad. Occurring in the regime. The steepdrop of moment coefficient at stall in the unsweptwings is not observed in the sweptback wings. I. INTROUCTION Paper submitted 11/19/9; revised 3/8/1; accepted 3/11/1. Author for correspondence: ShunChang Yen ( scyen@mail.ntou.edu.tw). *epartment of Mechanical and Mechatronic Engineering, National Taiwan Ocean University, Keelung, Taiwan, R.O.C. Many physical phenomena, such as, reattachment,, vortex, etc., evolve on the wing suction surface. The aerodynamic performance is closely related to the boundary flow patterns on wing surfaces. Fig. 1 depicts the vortex shedding behavior behind a sweptback wing and the coordinate system used in this study. The generally extends a large portion of wing surface and significantly changes the pressure distribution. Consequently, the aerodynamic performance is significantly changed. Mueller et al. [6, 7] experimentally studied the hysteresis loop in the curve of lift coefficient at low Reynolds number (Re) using a issaman 7769, Miley M613128, and NACA airfoils. Huang et al. [5] studied the aerodynamic performance by changing the surfaceflow mode on a NACA 12 airfoil. They found that the highest slope of lift coefficient (C ) occurs in the laminar regime and the increase rate of C decreases in the regime. In addition, the drag coefficient (C ) slightly decreases in the laminar regime, remains almost a constant in the regime and increases in the transition regime. Furthermore, the stall occurs in the turbulent regime. The stable vortex shedding behind a sweptback wing is initialized by a complex vortex on the wing surface and the unsteady flow behind the airfoil significantly affects the wing performances. Roshko [8] found that the ordinary Strouhal number (St) remained constant of about.21,.18, and.14 for a circular cylinder, 9 wedge and flat plate, respectively, in the range of 1 3 < Re < 1 5. The results indicated that the sharper the blockage body was, the lower the ordinary Strouhal number obtained. Zaman et al. [13] observed a low oscillation flow and found that the bluffbody shedding occurs at St.2 during the deep stall (α 18 ). However, at the onset of static stall (α 15 ), the Strouhal number is lower by an order than that in the deep stall. Huang and in [4] investtigated the vortex shedding and shearlayer instability on a NACA 12 wing. They revealed that the evolution of vortex shedding behind the airfoil at low angle of attack is closely related to the behavior of shearlayer instabilities. At high angles of attack, the low frequency shedding is superimposed by various high frequency shearlayer unstable waves. The characteristic modes laminar, subcritical, transition and supercritical modes of vortex shedding were determined by changing the Reynolds number and angle of attack. A systematic survey of surfaceflow patterns on a NACA 12 sweptback wing with Λ = 15 for < Re < was recently reported by Yen and Hsu [11]. Fig. 2 shows the distribution of surfaceflow patterns obtained by Yen and Hsu. The boundary layer flow structures were visualized using the surface oilflow scheme. Six characteristic flow regimes laminar,, leading

2 S.C. Yen: Aerodynamic Performance and Shedding Characteristics on a SweptBack Wing 163 Pitot tube Test section Flow y x y Λ z x Fig. 1. Schematic diagram of flow behavior on a sweptback wing. Rec A B1 B2 A: aminar B: B1: B2: eadingedge C: : E: Bluffbody wake C E Fig. 2. istribution of surfaceflow regimes utilized in this study and Yen and Hsu [11]. edge,, turbulent and bluffbody wake were categorized and studied using various chord Reynolds numbers and angles of attack. Furthermore, the distribution of characteristic surfaceflow modes are closely related to the configurations of vortex shedding behind the sweptback wing [11]. However, the properties of characteristic flow patterns and their effects on the aerodynamic performance were still not reported. In this study, the experimental results reveal the characteristic behaviors of surfaceflow modes and show the effects on the aerodynamic performances and unsteady flow structures behind the sweptback wings. The objectives of this research are (1) to measure the aerodynamic coefficients by using a six component balance, (2) to study the variation of moment coefficients between the unswept and sweptback wings, and (3) to measure Rotation and support mechanism Hotwire anemometer Force/Moment sensor system High speed data acquisition system Fig. 3. Experimental setup. FFT analyzer the vortex shedding frequency behind the sweptback wing using a hotwire anemometer. II. EXPERIMENTA SETUP Fig. 3 shows the schematic diagram of a closedreturn wind tunnel used to conduct the experiments. The test section is 6 cm 6 cm 12 cm in height, width and depth, respectively. A polished aluminumalloy plate was utilized as the testsection floor and three highly transparent acrylic panels were utilized for photography and visualization. The freestream velocity (u ) was measured using a Pitotstatic tube. Fig. 4 shows the profile of turbulent intensity (T.I.) against u. The maximum T.I. is <.2% for.56 < u < 4 m/s. In addition, the nonuniformity of the average velocity across the crosssection is <.5%. Fig. 4 displays the distribution of static pressure (P st P atm ) as a function of u. In addition, a aluminum plate with sharp leading and trailing edges was placed 5 mm over the testsection floor for controlling the boundary layer thickness. The thicknesses of boundary layer were about 4.3 mm and 1.65 mm [9] at u = 5. m/s and u = 3 m/s, respectively. The material of wing model is stainless steel and the wing airfoil is NACA 12 [1]. The sweepback angle (Λ) is 15 used in this study. Furthermore, the chord length is 6 mm and the wing span is 3 mm which yields a full span wing aspect ratio of 1. The wing model was mounted on a support and then inset through both the testsection floor and the boundarylayer thickness controlling plate. The vortexshedding frequency behind the sweptback wing was detected by a TSI 121T1.5 hotwire anemometer. The wire diameter and length are 5 μm and 1.5 mm to ensure the dynamic response frequency ranging from 15 to 25 khz. The hotwire signals were fed simultaneously to an FFT analyzer and a high speed PCbased data acquisition system. The data acquisition system embeds a sampleandhold function for removing the phaselag in the multichannel acquisition. The sampling rate and the elapsed time were set at 16, samples/sec and 2 seconds, respectively. The aerodynamic loadings were measured using a JR 3 Universal ForceMoment

3 164 Journal of Marine Science and Technology, Vol. 19, No. 2 (211) PstPatm (Pa) T.I. (%) P atm = mmhg T = K record length = 32 samples elapse time = 1.6 second sample rate = 2 khz f (Hz) Std Aperiodic Aperiodic Re c Re c St d =.8 α u (m/s) 3 4 Fig. 4. istributions of turbulent intensity (T.I.) and static pressure (P st P atm ) against the free stream velocity (u ) Fig. 5. istributions of vortexshedding frequency (f) and projected Strouhal number (St d ) versus angle of attack (α). System. The assembly of wing model and the JR 3 Universal ForceMoment System was mounted on a rotary supporter. The resolution of this rotary supporter is.12 degree/div. The JR 3 balance is a six degreeoffreedom force sensor and the output electronic signals are recorded by using a PCbased highspeed data acquisition system. The accuracy of u was affected primarily by the alignment of Pitot tube and pressure transducer. The uncertainty of u was 3% when a synchronized micro pressure calibration system was used and the Pitot tube was aligned carefully. The accuracy of α was controlled <.5% and the accuracy of vortexshedding frequency depends on the recording period of the hotwire anemometer and the sampling rate of FFT analyzer. Therefore, the accuracies of vortexshedding frequency, lift coefficient and drag coefficient were about ±.75%, ±1.5% and ±2.%, respectively. III. RESUTS AN ISCUSSION 1. Vortex shedding The vortexshedding frequency (f ) behind the sweptback wing was measured using a hotwire probe and the output signals of hotwire were recorded in an FFT analyzer. To remove the effects of wing junction and wingtip, the hotwire probe was installed at y/c = 2.5 which is on the center section of wing span, where y is the axis in the spanwise direction. In addition, the collected signals in the x direction (along the freestream direction) presented the similar frequency profiles. Consequently, the hotwire probe was installed at 1 < x/c < 5 to obtain the clear hotwire signals. The vortexshedding frequency varied with u was normalized using the nondimensional parameter projected Strouhal number (St d = f d/u ), where d is the projected chord length along the freestream direction. Fig. 5 plots the distributions of f versus α and St d versus α. Fig. 5 and shows an aperiodic region occurs in the transitional regime (7 < α < 22 ) [12]. Moreover, in Fig. 5, the vortexshedding frequency decreases with increasing α at various Re c. Fig. 5 delineates that the maximum St d of.51 occurs at α =. For α > 22, the St d is not changed with Re c and the relationship between St d and α is regressed as follows. St d =.8 α +.29, for 22 < α < 9, (1) where the unit of α is in degree.

4 S.C. Yen: Aerodynamic Performance and Shedding Characteristics on a SweptBack Wing Rτ C eadingedge τ (sec).3 C lag/c Re c Fig. 6. istributions of autocorrelation coefficient (R τ ) versus agrangian integral time scale (τ) and Taylor s integral length scale (l ag /C) versus angle of attack (α). Fig. 6 depicts the autocorrelation coefficient (R τ ) against the agrangian integral time scale (τ). The agrangian integral time scale is determined from the autocorrelation data and timeaveraged velocities by utilizing the Taylor s frozen flow hypothesis [1]. In addition, τ can be used to estimate the Taylor s integral length scale (l ag ) of shedding vortices and turbulent fluctuations. Fig. 6 indicates that τ is Furthermore, Fig. 6 shows the distribution of the normalized Taylor s integral length scale (l ag /C) against α at various chord Reynolds numbers (Re c ). Fig. 6 indicates that l ag /C is positively proportional to α and the effect of Re c on l ag /C is weak. 2. Aerodynamic performances Fig. 7 shows the distributions of lift coefficient, drag coefficient and pitching moment coefficient (C M ) about quarter chord length against α at Re c = 1 5. Fig. 7 shows that C increases monotonically with α in the and leadingedge regimes. The maximum C and the 6 CM (c) Sweptback wing Unswept wing Fig. 7. istributions of lift coefficient (C ), drag coefficient (C ) and (c) moment coefficient (C M ) against angle of attack (α) at Re c = 1 5. increase rate of C with respect to α (d(c )/dα) are 1.24 and 1.52 π/rad, respectively, occurring in the leadingedge regime. The theoretical value of d(c )/dα obtained from the analytical analysis of a twodimensional, thin, symmetric, and flatplate airfoil in inviscid flow is 2 π/rad [2]. Abbott and van oenhoff [1] indicated that (d(c )/dα is about 2.18 π/rad for a unswept NACA12 wing tested in the inviscid flow. Moreover, Fig. 7 shows that C decreases when the surface flow is transited into the regime. In the regime, the reattached turbulent surface flow separates, and therefore the second occurs. The second line moves toward the leading edge with increasing α. The minimum C of.93 occurs in the regime, and then C increases slightly as α is further increased into the turbulent regime. The liftrise phenomenon is induced from both the scavenging effect on the suction surface and the impact pressure on the pressure surface (Hoerner and Borst [3]).

5 166 Journal of Marine Science and Technology, Vol. 19, No. 2 (211) Moment axis Flow α = 1 α = 7 α = 9 eadingedge α = 14 α = 2 1/4C M y y Λ = 15 1/4C x + Sweptback wing model x 1/4C Aerodynamic center Fig. 8. Schematic diagram of lift and drag resultants and pitchmoment. Fig. 7 shows the distribution of C against α. The moves toward the leading edge and the length is shrunk with increasing α. The reduction of skin friction then competes with the increase of form drag. Consequently, the C does not change significantly in the regime. In the leadingedge regime, the increase of skin friction and the decrease of length lead to a discontinuous rising of C. In the regime, the reattached turbulent surface flow conducts a high skin friction on the suction surface. Consequently, the maximum increase rate of C with respect to α (d(c )/dα) of.49 π/rad occurs in the regime. The C increases almost linearly with the increase of α in the turbulent regime due to the significant increase of form drag. Fig. 7(c) shows that C M decreases as α increases for α < 1. In the and leadingedge regimes, the C M increases clockwise with α due to the wing stall. The local maximum C M occurs in the regime due to the increase of form drag. In the turbulent regime, the C M increases clockwise with α in consequence of a large increase of form drag. However, for a unswept wing, the C M decreases almost linearly with the increase of α for α < 2 while C M increases with α for 2 < α < 9 due to the pressure center moving toward the quarter chord point. Furthermore, the lift decreases and the pressure center moves toward the trailing edge for α > 9. Therefore, C M increases with α. In addition, Fig. 7(c) shows that the steepdrop of C M for a unswept wing is not observed in the sweptback wing. Fig. 8 schematically plots the lift () and drag () while the moments induced from the and are also displayed. In the regime ( < α < 7.5 ), the reaction point of moves toward the leadingedge, and both the and generate a counterclockwise moment. Consequently, C M increases with α. In the leadingedge regime (7.5 < α < 1 ), the reaction point of moves toward the leadingedge and passes the aerodynamic center. Therefore, leads a clockwise moment while produces a counterclockwise one. Consequently, the increase rate of C M is lowered. In the mode (1 < α < 16 ), the occurrence of stall conducts a sudden loss of lift and the reaction point of lift moves backward the trailingedge. Consequently, a sudden transition of C M curve occurs. In the turbulent mode (16 < α < 26 ), generates a counterclockwise moment with respect to the aerodynamic center. In addition, increases with α, and therefore leads a counterclockwise moment. Consequently, the sum of counterclockwise moment induced from and leads a increase in C M. Fig. 9 shows the distribution of liftdrag ratio (C /C ) as a function of α at Re c = 1 5. In the regime, the C /C increases from to 6.6 as α increases and the maximum C /C occurs at α = 7.5. In the leadingedge regime, the on the suction surface retards the increase rate of C /C, and therefore the C /C drops from 6.3 to 5.7 while α changes from 1 to In the and turbulent regimes, C /C decreases with the increase of α since C increases slightly and C increases rapidly in these two regimes. Fig. 9 shows the relationship between C and C at Re c = 1 5. In the regime, C does not change significantly with C. However, C increases gradually with C in the leadingedge regime. In regime, C increases while C decreases. In the turbulent regime, C changes significantly while C fixes approximately a constant. IV. CONCUSIONS The aerodynamic performances and vortex shedding of a finite sweptback wing were experimentally studied using different angles of attack and chord Reynolds numbers. The following conclusions are drawn from the results and discussion.

6 S.C. Yen: Aerodynamic Performance and Shedding Characteristics on a SweptBack Wing 167 C C/C Re c = eadingedge eadingedge C Fig. 9. istributions of liftdrag ratio (C /C ) versus angle of attack (α) and C versus C. (1) The relationship between St d and α is regressed as: St d =.8 α +.29, for 22 < α < 9. (2) C increases with α in the and leadingedge regimes; and the maximum increase rate of C with respect to α is 1.52 π/rad occurring in the leadingedge regime. (3) The maximum increase rate of C with respect to α is.49 π/rad occurring in the regime. (4) The steepdrop of C M occurring at stalling point for a unswept wing is not observed by utilizing the sweptback wing. ACKNOWEGMENTS This research was supported by the National Science Council of the Republic of China, under Grant No. NSC E196. NOMENCATURE B wing span (= 3 cm) C chord length (= 6 cm) C lift coefficient (= /qbc) C drag coefficient (= /qbc) C M moment coefficient about quarter chord point (= M/qbC) drag projected chord length along the freestream direction F vortexshedding frequency (Hz) lift M moment about quarter chord point Q dynamic pressure of free stream (= ρu 2 /2) Re c chord Reynolds number (= u C/ν) St d projected Strouhal number (= f d/u ) u free stream velocity Λ sweepback angle α angle of attack ρ density of air ν kinetic viscosity of air REFERENCES 1. Abbott, I. H. and von oenhoff, A. E., Theory of Wing Section, over Publications, New York, pp. 553 (1959). 2. Bertin, J. J. and Smith, M.., Aerodynamics for Engineers, Prentice Hall, Englewood Cliffs, New Jersey, pp (1989). 3. Hoerner, S. F. and Borst, H. V., Fluid ynamic ift, Mrs. iselotte A. Hoerner, Brick Town, New Jersey, (1975). 4. Huang, R. F. and in, C.., Vortex shedding and shearlayer instability of wing at lowreynolds numbers, AIAA Journal, Vol. 33, No. 8, pp (1995). 5. Huang, R. F., Shy, W. W., in, S. W., and Hsiao, F.B., Influence of surface flow on aerodynamic loads of a cantilever wing, AIAA Journal, Vol. 34, No. 3, pp (1996). 6. Mueller, T. J., The influence of laminar and transition on the low Reynolds number airfoil hysteresis, Journal of Aircraft, Vol. 22, No. 9, pp (1985). 7. Mueller, T. J., Pohlen,. J., Conigliaro, P. E., and Jansen, B. J., The influence of freestream dimensional on low Reynolds number airfoil experiments, Experiments in Fluids, Vol. 1, pp. 314 (1983). 8. Roshko, A., On the wake and drag of bluff bodies, Journal of the Aerospace Science, Vol. 22, pp (1955). 9. Shames, I. H., Mechanics of Fluid, 3rd ed., McGrawHill, Inc., Singapore, p. 632 (1992). 1. Tennekes, H. and umley, J.., A First Course in Turbulence, MIT Press, Cambridge, pp. 813 (1972). 11. Yen, S. C. and Hsu, C. M., Influence of boundary layer behavior on aerodynamic coefficients of a sweptback wing, ASME Journal of Fluids Engineering, Vol. 129, No. 6, pp (27). 12. Yen, S. C. and Hsu, C. M., Investigation on vortex shedding of a sweptback wing, Experimental Thermal and Fluid Science, Vol. 31, No. 8, pp (27). 13. Zaman, K. B. M. Q., McKinzie,. J., and Rumsey, C.., A natural lowfrequency oscillation of the flow over an airfoil near stalling conditions, Journal of Fluid Mechanics, Vol. 22, pp (1989).