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Aircraft Aerodynamics - Aerofoil Characteristics, 2D aerofoil vs 3D finite…
Aircraft Aerodynamics - Aerofoil Characteristics
NACA aerofoil
4 -digit series
NACA
4
Camber, 100*cmax/c
4
Location of ymax of camber 10* xcamber/c
1
Thickness, 100*t/c (% of chord)
5
Nomenclature
Camber line
zero camber = symmetric aerofoil
describes the curviness of the aerofoil
Angle of attack
angle between the inflow direction and the chord line
Thickness distribution
a related term is max thickness and its position
Distance between the upper and the lower surface
Chord line
the line on the aerofoil joining the leading edge and the trailing edge (most forward and most aft point)
5 digit series
NACA
2
Design cl is this first digit *3/20
3
Location of max camber (2nd & 3rd digit/2)
0
1
Thickness (% of chord)
2
An aerofoil is the cross-section of a wing, assuming no flow in the spanwise direction
2D aerofoil
lift drag moment act on a unit span
S = c x 1!
L', D' and M'
Units for L' and D' Newtons per metre, M' units are Newtons per metre times meter (or just Newton)
cl = L'/(0.5
rho_inf
(V_inf)^2*c)
cl = L'/(q_inf*c)
cd = D'/(0.5
rho_inf
(V_inf)^2*c)
cd = D'/(q_inf*c)
cm = M'/(0.5
rho_inf
(V_inf)^2*c^2)
cm = M'/(q_inf*c^2)
q_inf = 0.5
rho_inf
(V_inf)^2
Divide L' and D' by q_inf*c (normalise)
cl = cn
cos(alpha)-ca
sin(alpha)
cd = cn
cos(alpha)+ca
sin(alpha)
cn = N'/(q_inf*c)
ca= A'/(q_inf*c)
3D finite wing
Lift drag moment act on the whole wing
They are denoted as L, D and M
notice no accent as it is not per unit span
Newton
CL = L/(0.5
rho_inf
(V_inf)^2*S)
CD = D/(0.5
rho_inf
(V_inf)^2*S)
CM = M/(0.5
rho_inf
(V_inf)^2*c_hat)
History
1884, first aerofoil shape patented
Very thin aerofoils designed by Horatio Phillips
1903, first flight of the wright brothers
Used in wright flier still thin
1917, development oif the thick aerofoil
Gottingen 298 thick aerofoil section, fokker dr-1 (1917) aeroplane use the new aerofoil
internal wing structure (reduced drag from lack of wires)
can store things inside the wing
Higher max Cl
high rate of climb
enhanced manoeverability
1930's, systematic testing and development of aerofoil shapes by NACA
1965, supercritical aerofoils
a reduction in wave drag, by delaying the onset of shock, allowing for higher subsonic cruise speeds
Aerodynamic Forces
How is lift created?
Produced by a moving fluid
Momentum varies around the aerofoil
Drives a change in pressure
Imbalance in pressure around the upper and lower surface generates a net fluid force
Component of the fluid force normal to the flow is lift
Most of the pressure force is produced on the upper surface
Component of the fluid force parallel to the flow is drag
Pressure distribution, p
Acts normal to the surface
Shear stress distribution
tau
Acts tangentially along a surface
Integrate p and tau to find resultant aerodynamics force and moment
resolve into ^
Lift and Drag
Normal and Axial
L = Ncos(alpha)-Asin(alpha)
D = Nsin(alpha) + Acos(alpha)
references points
25% chord
leading edge
2D aerofoil vs 3D finite wing
