Stability 

Main Wing Stability Earlier we learned that if the wing is behind the c of g it is stable. Now we can be more precise and say that if the ac of the wing is behind the c of g the wing is stable. Use the movie below to clarify the concept. If the ac is ahead of the c of g the wing will be unstable. This concept is explained in the movie below. Stability Summary When the aerodynamic center is behind the c of g the wing is longitudinally stable. When the aerodynamic center is ahead of the c of g the wing is unstable. If the ac is exactly at the c of g the wing will have neutral longitudinal static stability. Pitching Moment Vs. Angle of Attack To facilitate a graphical analysis of the stability on an aircraft we will use Coefficient of Pitching Moment vs. Angle of Attack graphs. Pitching moment is calculated using an equation which is very similar to the lift or drag equations we have used earlier. However, keep in mind that the earlier equations generated units of force. Moment on the other hand has units of force times distance. Therefore, it should not surprise you that the equation for pitching moment is: M = CM x S x ½pV2 x c (c = arm, CM = Coefficient of pitching moment.) By definition positive values of M are nose up moments, and negative values are nose down moments. In the graph to the left we see that as CL increase CM decreases. In other words as angle of attack increases a nose down moment develops. This is a stable situation. As long as the CM vs. CL graph has a negative slope the wing is stable. The steeper the slope the more stable the wing. Developed by Geistware of Indiana© ., 1999. Updated January 1, 2003

Lateral  (pitch)  Motion around the lateral axis (diving on climbing)

Pitch control is provided by the elevators. Just as with roll and yaw described above a pitch damping effect quickly develops once the elevators are deflected. As a result the aircraft soon establishes a constant pitch rate once the elevators are deflected. Use the movie below to explore this effect.

Longitudinal (roll)  Motion around the longitudinal axis  (rolling)

 When aileron control is inputted one aileron moves up while the aileron on the opposite wing moves down resulting is an asymmetric lift between the wings. This causes the roll rate to increase away from the wing with the aileron in the down position and the aircraft will turn (or yaw) toward the wing with the aileron in the up position.

Vertical  motion around the normal axis  (Yaw)

Directional or yaw control is provided by the vertical fin and the rudder. This is probably the easiest of the three axis to visualize. 

aw damping is just like roll damping discussed above. 

When the rudder is first deflected the fin produces a net force, due to the camber created. This force, acting at an arm from the center of gravity creates a moment which starts the aircraft yawing, faster and faster.

But, as the aircraft yaws the angle of attack on the fin changes until the fin is at its zero lift angle of attack. Once this happens there is no longer a moment. From this point on the aircraft yaws at a constant rate.

It is worth noting that as soon as the pilot releases the rudder pedal the fin is no longer cambered. Therefore, it is not at the zero lift angle of attack anymore. A yawing moment will therefore, be created which will yaw the aircraft back in the opposite direction.

Of course since the fin is a symmetric airfoil the yaw rate will damp out at zero degrees of slip.