# Yaw Control

In most flight conditions it is desired to maintain the sideslip angle at zero. If the air­plane has positive yaw stiffness, and is truly symmetrical, then it will tend to fly in this condition. However, yawing moments may act upon the airplane as a result of unsymmetrical thrust (e. g., one engine inoperative), slipstream rotation, or the un­symmetrical flow field associated with turning flight. Under these circumstances, /3 can be kept zero only by the application of a control moment. The control that pro­vides this is the rudder. Another condition requiring the use of the rudder is the steady sideslip, a maneuver sometimes used, particularly with light aircraft, to in­crease the drag and hence the glide path angle. A major point of difference between the rudder and the elevator is that for the former trimming the airplane is a secondary and not a primary function. Apart from this difference, the treatment of the two con­trols is similar. From (3.9,3) and (3.9,6), the rate of change of yawing moment with rudder deflection is given by   (3.10,1) This derivative is sometimes called the “rudder power.” It must be large enough to make it possible to maintain zero sideslip under the most extreme conditions of asymmetric thrust and turning flight.

A second useful index of the rudder control is the steady sideslip angle that can be maintained by a given rudder angle. The total yawing moment during steady sideslip may be written Cn ~ Cnfi + Cns 8r

For steady motion, C„ = 0, and hence the desired ratio is £

8r

The rudder hinge moment and control force are also treated in a manner similar to that employed for the elevator. Let the rudder hinge-moment coefficient be given

by

Chr = bxaF + b28r (3.10,4)

The rudder pedal force will then be given by

P = G^V2FS^cr{b, aF + b28r)

= g| V2FSrcr[bx{-(3 + a) + b28r] (3.10,5)

where G is the rudder system gearing.

The effect of a free rudder on the directional stability is found by setting Chr = 0 in (3.10,4). Then the rudder floating angle is

(3-10,6)

The vertical-tail lift coefficient with rudder free is found from (3.9,3) to be

b

C’lf = aFaF – a —aF

b2

t ar bt

= aFaF 1-— тЧ (3-10,7)

aF b2)

The free control factor for the rudder is thus seen to be of the same form as that for the elevator (see Sec. 2.6) and to have a similar effect.