AE 451 Aeronautical Engineering Design I Aerodynamics. Prof. Dr. Serkan Özgen Dept. Aerospace Engineering December 2017
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1 AE 451 Aeronautical Engineering Design I Aerodynamics Prof. Dr. Serkan Özgen Dept. Aerospace Engineering December 2017
2 Lift curve 2
3 Lift curve slope 3
4 Subsonic lift curve slope C Lα = AR2 β 2 η 2 2πAR 1 + tan2 Λ max,t β 2 S exposed S F Valid until M dd, fairly accurate until M=1. β 2 = 1 M 2 η: airfoil efficiency, = 0.95 for most airfoils. F = d/b 2, fuselage lift factor. AR eff = AR( h b); effective AR with endplates, h: height of the endplate. AR eff 1.2AR; effective AR with winglets. 4
5 Supersonic lift curve slope Theory: C Lα = 4 β Practice: use the charts valid for trapezoidal wings. Correct the values read with S exposed S F 5
6 Supersonic lift curve slope 6
7 Maximum lift (clean) For moderate to high aspect ratio wings with moderate sweep and high leading edge radius: C L,max = 0.9c l,max cosλ c/4 If a wing has low AR or high sweep and a sharp leading edge, maximum lift will increase due to leading edge vortices. This is a function of the shape of the upper surface of the leading edge: y = y 0.06c y 0.015c 7
8 Maximum lift (clean) 8
9 Maximum lift (clean) 9
10 Maximum lift (clean) For high aspect ratio wings: C L,max = c l,max C L,max c l,max + C L,max, Correct for F, S exposed. S α CL,max = C L,max C L,α + α 0L + α CL,max 10
11 Maximum lift (clean) 11
12 Maximum lift (clean) 12
13 Maximum lift (clean) 13
14 Maximum lift (clean) A wing has low AR if: AR 3 C 1 +1 cos Λ LE. C L,max = C L,max,base + C L,max. α CL,max = α CL,max,base + α CL,max 14
15 Maximum lift (clean) 15
16 Maximum lift (clean) 16
17 Maximum lift (clean) 17
18 Maximum lift (clean) 18
19 Maximum lift (clean) At transonic speeds, maximum lift is limited by structural buffeting and controllability considerations rather than aerodynamics. 19
20 Maximum lift (with high lift devices) 20
21 Maximum lift (with high lift devices) 21
22 Maximum lift (with high lift devices) 22
23 Maximum lift (with high lift devices) Trailing edge devices decrease the stall angle of attack by increasing the pressure drop over the top of the airfoil promoting flow separation. In order to increase α stall a leading edge device must be used. C L,max = 0.9 c l,max S flapped S cosλ HL α 0L = α 0L,airfoil S flapped S cosλ HL HL: hinge line of the high lift device 23
24 Maximum lift (with high lift devices) For takeoff, the increments of about 60-80% of the increment calculated above should be used. Maximum lift occurs at a flap setting of about 40 o -45 o. α0l,airfoil 15 o (landing setting), α0l,airfoil 10 o (takeoff setting), 24
25 Maximum lift (with high lift devices) 25
26 Maximum lift (with high lift devices) 26
27 Maximum lift (with high lift devices) Leading edge devices increase lift by: Increasing camber, Increasing wing area, Delaying separation. Leading edge devices are particularly useful at high α. During takeoff and landing, they are useful when in combination with trailing edge devices as they prevent stall. 27
28 Estimation of C Do, equivalent skin friction method C Do = S wet S C fe C fe : equivalent skin friction coefficient is a function of the Reynolds number, Re. 28
29 Equivalent skin friction coefficients 29
30 Wetted area ratio 30
31 Estimation of C Do, component build-up method Total parasite drag coefficient: ) C Do subsonic = C fcff c Q c S wet,c S + C D,misc + C D,L&P C fc : flat plate skin friction coefficient, C fc = C fc Re, M, k ; k: skin roughness. FF c : form factor, estimates pressure drag due to separation, Q: interference factor. C D,misc : drag of flaps, landing gears, upswept aft fuselage, base area. C D,L&P : drag of leakages and protuberances. 31
32 Flat plate skin friction coefficient Laminar flow: C f = Re, Re = ρ V l, μ l: characteristic length. Turbulent flow: C f = log Re M
33 Flat plate skin friction coefficient 33
34 Flat plate skin friction coefficient If the surface is rough, the skin friction coefficient will be higher. The smaller of the cut-off Reynolds number and the actual Reynolds number shall be used. Subsonic flow: Re cutoff = l k 1.053, Transonic or Supersonic flow: Re cutoff = l k M
35 Flat plate skin friction coefficient 35
36 Component form factors Wing, tail, strut and pylon: FF = x/c m t c t c M 0.18 cos Λ m 0.28 x c m : chordwise location of the maximum thickness point, Λ m : sweep angle at the same location 36
37 Component form factors Fuselage and smooth canopy: f = l d = FF = f 3 + l 4 π A max : fineness ratio. Nacelle and external stores: FF = f f
38 Component form factors For a tail surface with a hinged control surface: +10% A square sided fuselage: +40% For a two piece canopy: +40% For an external boundary-layer diverter for a fuselage mounted inlet: Double wedge: FF = 1 + d l, Single wedge: FF = 1 + 2d l. 38
39 Component form factors 39
40 Component interference factors Nacelle or external store mounted on wing or fuselage: Q=1.5. Nacelle or external store mounted on wing or fuselage: Q=1.3 (if mounted less than one diameter away). Nacelle or external store mounted on wing or fuselage: Q=1.1 (if mounted more than one diameter away). Wingtip mounted missiles: Q=1.25. High-wing, mid-wing or a well-filleted low-wing: Q=1.0. Unfilleted low-wing: Q= Fuselage: Q=1.0. Tail surfaces: Q=1.03 (V-tail), 1.08 (H-Tail), (conventional tail). 40
41 Miscellaneous drag Upswept aft fuselage: D q = 3.83θ2.5 A max. 41
42 Miscellaneous drag Landing gear: summation of the drags of the wheels, struts, and other gear components. Q=1.2, *1.07 for retractable landing gears accounting for the hollow landing gear well. 42
43 Miscellaneous drag Flaps: C Do,flap = F flap c flap c S flapped S δ flap 10 o. F flap = : plain flaps, F flap = : slotted flaps. Speed brakes: Fuselage mounted: D q = 1.0A frontal Wing mounted: D q = 1.6A frontal 43
44 Miscellaneous drag Canopies (transport and light aircraft): D q = 0.50A frontal,wind shield Cannon port: D q = 0.2 ft2. 44
45 Leakage and protuberance drag Antennas, lights, door edges, fuel vents, control surface external hinges, actuator fairings, rivets, rough or misaligned panels Jet transports and bombers: 2-5% parasite drag, Propeller aircraft: 5-10%, Fighters: 10-15% (old), 5-10% (new). 45
46 Supersonic Wave Drag For supersonic skin friction drag Q = FF = 1. ) C D0 s.sonic = C fcs wet S + C D,misc + C D,L&P + C D,wave Leakage and protuberance drag percentages apply only to skin-friction drag. For preliminary wave drag analysis (M 1.2): 0.77 D = E q wd (M 1.2) πλ LE,deg wave 100 D q Sears Haack 46
47 Sears-Haack body 47
48 Supersonic wave drag E wd : wave drag efficiency factor. =1.0 for a perfect Sears-Haack body, =1.2 for a smooth volume distribution, blended delta wing, = for a supersonic fighter, bomber. D q Sears Haack = 9π 2 A max l 2 ; subtract inlet capture area. l: aircraft length length with constant cross sectional area. 48
49 Transonic wave drag Boeing formulation: M DD = M DD,L=0 LF DD 0.05C L,design wing. 49
50 Transonic wave drag 50
51 Transonic wave drag 51
52 Transonic wave drag L n : length of fuselage from nose to the location where fuselage cross section becomes constant. d: equivalent diameter of the fuselage there. Choose the smaller of the M dd found for wing and fuselage for the drag divergence Mach number of the airplane. 52
53 Transonic wave drag For initial analysis: 53
54 Transonic wave drag M 1.2: use supersonic wave drag expression. C D,wave M = 1.05 = C D,wave M = 1.2. C D,wave M = 1.0 = C D,wave M= M cr = M DD C D M DD = C D M cr
55 Complete drag build-up Subsonic drag: skin friction drag (including form factor and interference) + miscellaneous drag + leakage & protuberance drag Supersonic drag: skin friction drag + miscellaneous drag + leakage and protuberance drag + wave drag. 55
56 Complete drag build-up 56
57 Complete drag build-up 57
58 Drag due to lift (induced drag) Induced drag coefficient: K = 1 πare Straight-winged airplane: e = AR (Λ LE < 30 o ) Swept winged airplane: e = AR 0.68 cos Λ 0.15 LE 3.1 (Λ LE > 30 o ) At supersonic speeds: K = AR(M2 1) 4AR M cos Λ LE 58
59 Drag due to lift (induced drag) Flap effect on induced drag: C Di = K f 2 ( C L,flap ) 2 cos Λ c 4, K f = 0.14: full span flaps, K f = 0.28: partial span flaps. 59
AE 451 Aeronautical Engineering Design I Aerodynamics. Prof. Dr. Serkan Özgen Dept. Aerospace Engineering December 2015
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