Lecture # 08: Boundary Layer Flows and Drag

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1 AerE 311L & AerE343L Lecture Notes Lecture # 8: Boundary Layer Flows and Drag Dr. Hui H Hu Department of Aerospace Engineering Iowa State University Ames, Iowa 511, U.S.A

2 y AerE343L #4: Hot wire measurements in the wake of an airfoil Pressure rake with 41 total pressure probes (the distance between the probes d=2mm) x 8 mm Lab#3 Test conditions: Velocity: V=15 m/s Angle of attack: AOA=, and 12 deg. Date sampling rate: f=1hz Number of samples: 1, (1s in time) No. of points: 2~25 points Gap between points: ~.2 inches Lab#4 Hotwire probe

AerE343L #4: Hot wire measurements in the wake of an airfoil 6 4 2 Y /C *1 Lab#4 Hotwire probe -2-4 -6 25 m/s shadow region vort: -4.5-3.5-2.5-1.5 -.5.5 1.5 2.5 3.5 4.

3 AerE343L #4: Hot wire measurements in the wake of an airfoil Y /C *1 Lab#4 Hotwire probe m/s shadow region vort: X/C *1 FFT arbitary scale Force -Z component (N) time sequence with data sampling rate of 1 Hz Freqency (Hz)

4 AerE343L #4: Hot wire measurements in the wake of an airfoil Lab#4 Hotwire probe Required data for the lab report: 1. Wake velocity profiles at AOA = and 12 deg 2. Wake turbulence intensity profiles at AOA = and 12 deg. 3. Estimated drag coefficients at AOA=, and 12 deg. 4. FFT transformation to find vortex shedding frequency in the wake of the airfoil 5. Discussions based on the measurement results

5 Boundary Layer Flows Y X τ w U = μ y wall Which one will induce more drag? Laminar boundary layer? Turbulent boundary layer?

The major and most important difference between the two types of airfoil is this, the thickest part of a laminar wing occurs at 5% chord while in the conventional design the thickest part

6 CONVENTIONAL AIRFOILS and LAMINAR FLOW AIRFOILS Laminar flow airfoils are usually thinner than the conventional airfoil. The leading edge is more pointed and its upper and lower surfaces are nearly symmetrical. The major and most important difference between the two types of airfoil is this, the thickest part of a laminar wing occurs at 5% chord while in the conventional design the thickest part is at 25% chord. Drag is considerably reduced since the laminar airfoil takes less energy to slide through the air. Extensive laminar flow is usually only experienced over a very small range of angles-of-attack, on the order of 4 to 6 degrees. Once you break out of that optimal angle range, the drag increases by as much as 4% depending on the airfoil

7 Flow Separation

8 Aerodynamic Performance of An Airfoil Lift Coefficient, C l L = 1 ρv 2 C l 2 c C L =2πα Experimental data Airfoil stall Y /C * m/s shadow region Before stall -6 vort: Angle of Attack (degrees) X/C * Drag Coefficient, C d D = 1 ρv 2 C d 2 c Experimental data Y /C * m/s shadow region After stall.5 Airfoil stall -4 vort: Angle of Attack (degrees) X/C *1

Flow Separation and Transition on Low-Reynolds Reynolds-number number Airfoils Low-Reynolds Reynolds-number number airfoil (with Re<5,) aerodynamics is important for both military and civilian

9 Flow Separation and Transition on Low-Reynolds Reynolds-number number Airfoils Low-Reynolds Reynolds-number number airfoil (with Re<5,) aerodynamics is important for both military and civilian applications,, such as propellers, sailplanes, ultra-light light man- carrying/man-powered aircraft, high-altitude vehicles, wind turbines, unmanned aerial vehicles (UAVs( UAVs) ) and Micro-Air Air-Vehicles (MAVs). Since laminar boundary layers are unable to withstand any significant adverse pressure gradient,, laminar flow separation is usually found on low-reynolds Reynolds-number number airfoils. Post- separation behavior of the laminar boundary layers would affect the aerodynamic performances of the low-reynolds Reynolds-number number airfoils significantly Separation bubbles are usually found to form on the upper surfaces of low-reynolds Reynolds-number number airfoils. Separation bubble would burst suddenly to cause airfoil stall at high AOA when the adverse pressure gradient becoming too big. C L Thin airfoil theory C L (Re=68,) C D (Re=68,) Copyright by Dr. Hui Iowa State University. All angle Rights of attack Reserved! (degree) C D

Surface Pressure Coefficient distributions (Re=68,) Separation point C P -3. -2.5-

10 Surface Pressure Coefficient distributions (Re=68,) Separation point C P Turbulence transition Reattachment point AOA = 6 deg AOA = 8 deg AOA = 9 deg AOA = 1 deg AOA = 11 deg AOA = 12 deg AOA = 14 deg Y/C Typical surface pressure distribution when a laminar separation bubble is formed (Russell, 1979) X/C GA (W)-1 1 airfoil (also labeled as NASA LS(1)-417 ) X / C AOA (degree) 12. Transition 11.5 Reattachm ent Separation X/C

11 Laminar Separation Bubble on a Low-Reynolds Reynolds-number number Airfoil 1 1 Y/C*1 Y/C* Spawise vorticity GA (W)-1 airfoil 1 m/s X/C*1 1 m/s Instantaneous flow field Spanwise Vorticity (1/s * 1 3 ) X/C*1 Y/C*1 Y/C* separation U m/s: X/C*1 1 m/s PIV measurement results at AOA = 1 deg, Re=68, (Hu et al., ASME Journal of Fluid Engineering, 28) GA (W)-1 airfoil Ensemble-averaged flow field Reattachment Um/s: reattachment X/C*1

12 Stall Hysteresis Phenomena Stall hysteresis, a phenomenon where stall inception and stall recovery do not occur at the same angle of attack, has been found to be relatively common in low-reynolds-number airfoils. When stall hysteresis occurs, the coefficients of lift, drag, and moment of the airfoil are found to be multiplevalued rather than single-valued functions of the angle of attack. Stall hysteresis is of practical importance because it produces widely different values of lift coefficient and lift-to-drag ratio for a given airfoil at a given angle of attack. It could also affect the recovery from stall and/or spin flight conditions. Lift coefficient Increasing AOA decreasing AOA Lift coefficient Increasing AOA decreasing AOA AOA Angle of Attack Lift coefficient curve of a typical airfoil AOA Angle of Attack Lift coefficient curve with stall hysteresis

13 Measured airfoil lift and drag coefficient profiles Lift Coefficient, C l Hysteresis loop AOA increasing AOA decreasing Drag Coefficient, C d AOA increasing AOA decreasing Hysteresis loop Angle of Attack (degree) Angle of Attack (degree) GA(W)-1 1 airfoil, Re C = 16, The hysteresis loop was found to be clockwise in the lift coefficient profiles,, and counter-clockwise clockwise in the drag coefficient profiles. The aerodynamic hysteresis resulted in significant variations of lift coefficient, C l, and lift-to to-drag ratio, l/d, for the airfoil at a given angle of attack. The lift coefficient and lift-to to-drag ratio at AOA = 14. degrees were found to be C l = 1.33 and l/d = 23.5 when the angle is at the increasing angle branch of the hysteresis loop. The values were found to become C l =.8 and l/d = 3.66 for the same AOA=14. degrees when the angle is at the deceasing angle branch of the hysteresis loop

14 PIV Measurement results Y /C *1-2 Y /C *1 separation bubble m/s shadow region vort: m/s shadow region vort: X/C *1 Lift Coefficient, C l AOA decreasing AOA increasing X/C *1 Y /C * m/s shadow region Angle of Attack (degree) Y /C * m/s shadow region -4 vort: vort: X/C * X/C *1 Copyright (Hu, Yang, by Dr. Igarashi, Hui Hu Iowa State of Aircraft, University. Vol. All Rights 44. No. Reserved! 6, 27)

15 Refined PIV Measurement Results Y /C *1 1 5 Y /C *1 1 5 Y /C * m/s vort: X/C * Lift Coefficient, C l -1 2 AOA decreasing AOA increasing Angle of Attack (degree) Y /C * m/s X/C *1 25 m/s 25 m/s -5 vort: vort: X/C * X/C *1 Copyright (Hu, Yang, by Dr. Igarashi, Hui Hu Iowa State of Aircraft, University. Vol. All Rights 44. No. Reserved! 6, 27) -1

16 Aerodynamics of Golf Ball

2 smooth-ball rough-ball golf-ball -.4-2. -1.5-1. -.5.5 1. 1.5 2.

Smooth ball Rough ball Distance (X/D) Golf ball U m/s: -1-5 5

17 Laminar Flows and Turbulence Flows 1. Re=1, Centerline Velocity (U/U ) smooth-ball rough-ball golf-ball Smooth ball Rough ball Distance (X/D) Golf ball U m/s: U m/s: U m/s: Y/D Y/D Y/D X/D X/D 1 X/D -1

Lecture # 08: Boundary Layer Flows and Controls

Lecture # 08: Boundary Layer Flows and Controls AerE 344 Lecture Notes Lecture # 8: Boundary Layer Flows and Controls Dr. Hui Hu Department of Aerospace Engineering Iowa State University Ames, Iowa 511, U.S.A Flow Separation on an Airfoil Quantification