There are eight factors that affect lift. The first three are readily apparent: Density (ρ), velocity (V), and surface area (S). The five remaining factors are all accounted for within the coefficient of lift. As stated, both angle of attack (α) and camber affect the production of lift. The remaining three factors are not so easily discernable. They are aspect ratio (AR), viscosity (μ) and compressibility.
When an airfoil is exposed to greater dynamic pressure (q), it encounters more air particles and thus produces more lift. Therefore, lift is dependent upon the density of the air (i.e., the altitude) and the velocity of the airflow. An increase in density or velocity will increase lift.
Since lift is produced by pressure, which is force per unit area, it follows that a greater area produces a greater force. Therefore, an increase in wing surface area produces greater lift.
The pilot has no control over aspect ratio, viscosity and compressibility. Aspect ratio deals
with the shape of the wing and will be briefly discussed later. Viscosity affects the aerodynam- ic force since it decreases the velocity of the airflow immediately adjacent to the wing’s surface. Although we consider subsonic airflow to be incompressible, it does compress slightly when it encounters the wing. Because there is no way to control aspect ratio, viscosity, or compress- ibility, they will be ignored in this discussion unless specifically addressed.
The coefficient of lift depends essentially on the shape of the airfoil and the AOA. Flaps are the devices used to change the camber of an airfoil, and are used primarily for takeoffs and landings. When employed, they will be lowered to a particular setting and remain there until takeoff or landing is complete. This allows us to consider each separate camber situation (i.e.
flap setting) individually and plot CL against AOA. AOA is the most important factor in the coefficient of lift, and the easiest for the pilot to change.
These curves are for three different airfoils: One symmetric, one negative camber and one positive camber. The shape of the CL curve is similar for most airfoils. At zero angle of attack, the positive camber airfoil has a positive CL, and the negative camber airfoil has a negative CL. The point where the curves cross the horizontal axis is the AOA where the airfoil produces no lift
(CL = 0). At zero AOA the symmetric airfoil has CL = 0. The positive camber airfoil must be at a negative AOA, and the negative camber airfoil must be at a positive AOA for the CL to equal zero.
As angle of attack increases, the coefficient
of lift initially increases. In order to maintain
level flight while increasing angle of attack, velocity must decrease. Otherwise, lift will be greater than weight and the airplane will climb. Velocity and angle of attack are inversely related in level flight.
↓
L = 12 ρ V 2 S C L
As angle of attack continues to increase, the coefficient of lift increases up to a maximum value (CLmax). The AOA at which CLmax is reached is called CLmax AOA. Any increase in angle of attack beyond CLmax AOA causes a decrease in the coefficient of lift. Since CLmax is the greatest coefficient of lift that can be produced, we call CLmax AOA the most effective angle of attack. Note that as long as the shape of an airfoil remains constant, CLmax AOA will remain constant, regardless of weight, dynamic pressure, bank angle, etc.
