Airfoil selection - ITLiMS

Wing
Airfoil selection
• Aerodynamic characteristics (K max, C Lmax,
stall characteristics)
• Structural reasons;
1

Leading
edge
Camber line
Airfoil geometry
Airfoil geometry
Chord
Maximum
thickness
position Maximum camber
position
Maximum
camber
Maximum
thickness
Trailing
edge
2

V ∞
AoA=0°
AoA=10°
Angle of attack definition
AoA
AoA=15°
Stall
AoA=20°
Separation point
3

Airfoil aerodynamic characteristics
Lift coefficient (C z or C L)
CLMAX
α 0
/dα
dC l
Stall
≡ a
α KR
Drag coefficient (C x or C D )
Airfoil aerodynamic characteristics
Design C z
4

Airfoil aerodynamic characteristics
Gliding ratio (C z / C x)
Power factor
(C z 3 / Cx 2 lub Cz 1,5 /Cx)
Airfoil aerodynamic characteristics
Pitching moment coefficient C m
Derivative dCm/dCz
is an indicator of
stability.
It is negative for
stable aeroplanes
and positive for
unstable aeroplanes.
5

Maximum thickness – t/c
Chord - c
Maximum
thickness - t
Effect of airfoil thickness on lift coefficient
6%
8%
10%
12%
14%
16%
18%
20%
6

Effect of airfoil thickness on lift coefficient
Effect of airfoil thickness on drag coefficient
6%
8%
10%
12%
14%
16%
18%
20%
7

6%
8%
10%
12%
14%
16%
18%
20%
6%
8%
10%
12%
14%
16%
18%
20%
Effect of airfoil thickness on gliding ratio
Effect of airfoil thickness on power factor
8

Camber line
0%
0,5%
1%
1,5%
2%
2,5%
3%
3,5%
Camber
Maximum
camber
Effect of airfoil camber on lift coefficient
9

Effect of airfoil camber on drag coefficient
0%
0,5%
1%
1,5%
2%
2,5%
3%
3,5%
0%
0,5%
1%
1,5%
2%
2,5%
3%
3,5%
Effect of airfoil camber on gliding ratio
10

0%
0,5%
1%
1,5%
2%
2,5%
3%
3,5%
Effect of airfoil camber on power factor
Effect of airfoil camber on moment coefficient
0%
0,5%
1%
1,5%
2%
2,5%
3%
3,5%
11

Position of
maximum
thickness
Position of maximum thickness
Boundary layer development
laminar turbulent
transition
Maximum
thickness
separation
separated
12

Effect of airfoil „laminarity” on drag coefficient
15%
20%
25%
30%
35%
40%
45%
50%
Effect of airfoil „laminarity” on lift coefficient
15%
20%
25%
30%
35%
40%
45%
50%
13

15%
15%
Effect of camber line shape on moment coefficient
28%
22%
28%
22%
35%
0°
2°
4°
6°
Effect of camber line shape on gliding ratio
35%
0%
2%
4%
6%
14

Reynolds
number effect on
aerodynamic
coefficients
Effect of
Mach
number on
lift
coefficient
15

Effect of
Mach
number
on drag
coefficient
Effect of Mach number on moment
coefficient
16

Critical Mach
number
Historical values of an aeroplane airfoil thickness
as a function of Mach number
17

Critical Mach
number
Critical Mach
number
18

Calculate Reynolds number for design airspeed
Airfoil selection
Re>3 000 000 3 000 000>Re>500 000 500 000>Re
Calculate Mach number
for maximum airspeed
M max 0,75
Wortmann catalogue
„Stuttgarter
Profilkatalog” Vol.1 i 2
Selig catalogue
„Summary of low speed
airfoil data” Vol.1-3
„Airfoils at low speeds”
Abbot catalogue „Theory of the wing section”, raport NACA 824, NASA TN D-7428
Calculate Reynolds number for design
airspeed
Find characteristics for Re des i M des
Calculate C L for design airspeed
Supercritical airfoil eg. NASA SC(2) 714
NASA TM X-1109
NASA TM X-2977
NASA TP 2969
Airfoil
selection
Compare C D for C Ldes of available airfoils and select few best airfoils
Compare C Lmax of selected airfoils
Compare stall character of selected airfoils
Compare C M of selected airfoils
Select an airfoil with a combination of above features best suiting to the
aeroplane mission
19

Remaining wing features
• Wing incidence;
• Mean aerodynamic chord mac, c
• Wing area (reference area) S;
• Wing span b;
• Wing aspect ratio A;
• Wing dihedral;
• Wing sweep angle (leading edge ΛLE, quarter chord Λc/4); • Taper ratio λ;
• Geometrical and aerodynamic twist;
• Winglets
• Leading edges extensions;
Wing incidence angle
An angle between root chord and fuselage
longitudinal axis
20

c R
c R
c T
S
W
S =
W / S
b = A ⋅ S
c R
T
=
b/2
2⋅
S
[ b ⋅ ( 1 + λ)
]
c = λ ⋅ c
R
Taper ratio
c T
c
λ =
c
T
R
Straight wings:
λ=0.4÷0.5
Swept wings:
λ=0.2÷0.3
Mean aerodynamic chord mac, c
0,25mac
Y
c
c R
c T
⎛ 2 ⎞
c = ⎜⎝
⎟⎠
⋅ cROOT
⋅
3
⎛ ⎞
= ⎜ ⎟⋅
⎝ 6 ⎠
b
Y
2 ( 1+
λλλλ + λλλλ )
( 1+
λλλλ )
[ ( 1+
2⋅
λ )( 1+
λ)
];
;
21

Vortices generated by a wing
Vortices generated by a wing and effect
of aspect ratio on drag coefficient
A
=
b
S
2
C
D
=
C
D0
2
CL
+
π ⋅ A ⋅e
22

Effect of aspect ratio (A, AR) on lift
coefficient
C
b 1 b2
b 3
Wing dihedral angle
φ – an angle between
chords’ plane and
horizontal plane
b 4
L
α
Unswept
= C
l
α
⋅
⎛
⎜⎜
⎝
C
l
π
⎞
⎟⎟ +
⎠
A
⎛
⎜⎜
⎝
A
Wing dihedral
Subsonic swept
Supersonic swept
ϕ
α
C
Helmbolt equation
=
l
π
α
2
⎞
⎟⎟
⎠
b 3

Λ c/4
M eff =M ∞ cos(Λ LE )
M kryt ~1/cos m (Λ LE )
Wing sweep reduces
effective Mach number.
q eff=q ∞cos 2 (Λ LE)
W~tan 2 (Λ LE)
Wing sweep Λ LE , Λ c / 4,
Λ t / c
tan LE
c / 4
Λ LE
Line connecting quarter
chords along the wing span
[ ( 1 − λ)
/ A ⋅ ( + λ)
]
Λ = tan Λ +
1
Wing sweep
Λ t/c
Λ LE
M
McosΛ LE
24

Wing sweep effect on dC L /dα
dC
dα
L
=
2 +
4 +
β =
eff
( A ⋅β)
2⋅
π ⋅ A
2
2 ⎛ tan ( Λ ) ⎞ t / c
⋅ ⎜⎜ 1 + ⎟ 2
⎝ β ⎠
2
1− Meff
M M cos Λ
= ∞
Wing sweep effect on separation
LE
25

Wing sweep at supersonic speeds
Winglets
26

Geometric twist
Geometric twist
Wing twist
Aerodynamic twist
Wing twist
Aerodynamic twist
27

AoA
V
Leading
Edge
eXtensions
Delta wings
28

LEX effect on lift coefficient
Vortex generators
RAF Museum Hendon
29