AS309 Review Notes Unit II
Airfoil Section Terminology
 Upper and lower surfaces
 Leading and trailing edges
 Chord line
 Mean camber line
 Leading edge radius
 Maximum camber
 Maximum thickness
Airfoil Section Characteristics Affecting Lift
 Shape of mean camber line (more camber means more lift)
 Maximum thickness and location
 Leading edge radius
Lift on an airfoil develops when static pressure on the upper surface decreases as dynamic pressure
increases, in accordance with Bernoulli’s Principle.
Lift Terminology
 Flight path and relative wind
 AOA ()
 Aerodynamic Force (AF)
 Center of Pressure (CP) vs. Aerodynamic Center (AC)
 Laminar vs. turbulent flow
AF – components are lift (perpendicular to relative wind) and drag (parallel relative wind)
Total Drag is the scalar sum of Parasite Drag and Induced Drag. Induced drag results form wingtip
vortices. An infinite span wings produces parasite drag but no induced drag. We discuss finite span wings
in Unit III of AS309.
Boundary Layer – every surface of an aircraft in flight has a boundary layer of air surrounding it. This
layer begins at the skin surface and extends outward to the point where the movement of the aircraft no
longer disturbs the air surrounding the aircraft.
Parasite Drag Categories
 Skin Friction Drag – boundary layer friction due to air viscosity.
 Form Drag – air piling up in front of non-streamlined surfaces.
 Interference Drag – turbulence due to conflict between two boundary layers.
 Leakage Drag – caused by air flowing from high to low pressure, potentially interrupting
smooth boundary layer flow.
Drag Equation: D = CDqS = CDV2S / 295.37.
Non-symmetrical airfoil produces a pitching moment at zero lift; a symmetrical airfoil doesn’t.
CP moves considerably along the chord line as AOA or airspeed varies, creating changes in the airplane’s
pitching moment. The AC remains 23% – 27% of chord line aft of leading edge in subsonic flight.
(Moves to 50% for transonic / supersonic flight, as discussed in Unit III.) For a given airspeed, pitching
moment at AC is constant regardless of AOA.
Equilibrium Conditions: Fx = Fy = Fz = Mx = FMy = Mz = 0.
Rogers’s AS309 Review Notes, Unit II (revised 29 February 2000): Page 1
AOA Indicator
 Measures AOA (angle between relative wind and chord line)
 Calibrated in units as opposed to degrees to facilitate use in different types of airplanes
 Often used in conjunction with a stall indexer
 AOA for max range, max endurance, stall, &c. constant regardless of gross weight, angle of
bank, or altitude
Stall commences when the boundary layer begins to separate from the upper wing surface.
 Conventional (1G) vs. accelerated stall
 Third type is shock wave stall, which commences in transonic flight.
Boundary Layer Theory
 Reynolds number Re = V x / ) V is velocity in ft/sec; x is flow distance;  is kinematic
viscosity.
 High Re associated with turbulent flow; low Re associated with laminar flow
 Turbulent flow gives high skin friction drag but delayed boundary layer separation and lower
total drag.
 Turbulent flow on a sphere in airflow reduces wake size and total drag, even though skin
friction drag is higher.
 More lift (from increased possible AOA) and less total drag associated with high R e.
High Lift Devices (Energy Adders)
 Boundary layer energized to delay separation at high AOA.
 Vortex generators
 Boundary Layer (BL) control (can be passive or effected by high pressure air from jet engine
compressor section).
 Leading edge slots are example of passive BL control device.
 BL control devices delay separation of BL in adverse pressure gradient region (aft portion of
upper surface of wing).
Factors Affecting Lift
 Dynamic pressure q =  V2 / 2 (airspeed, density ratio )
 Planform area of wing
 Shape of airfoil section
 Air viscosity
 Compressibility effect
 AOA
Formulae





q =  V2 / 2 = 0 V2 / 2, since  = 0  (V in ft/sec)
L = CL q S = CL (0 / 2)  V2 S (V in ft/sec; S in ft2) = CLV2S / 295.37 (V in kts, i.e. nm/hr)
D = CD  V2 S / 295.37 (V in kts, i.e. nm/hr)
C L 1
V2
W2
1
cos 1
(V is TAS)




C L 2
V1
W1
2
cos  2
CL
L
GD
  GR 
(GR is glide ratio; GD is glide distance; AA is absolute altitude)
CD D
AA
Test Problems
 TAS changes as a function of weight, density altitude, coefficient of list, and bank angle (see
Quiz 5)
 Force on elevator necessary to correct nose pitch due to lift (as in class notes)
 Computing glide distance from glide ratio or (L/D) max, or (L/D)max from GD and GR
Rogers’s AS309 Review Notes, Unit II (revised 29 February 2000): Page 2