Wing Yawing Derivatives C/r, Cnr

The following methods are simplified versions of those given in USAF Datcom. They apply to rigid straight-tapered wings in subsonic flow at low values of CL.

The derivative C,:




—1 is the slope of the rolling moment due to yawing at zero lift given by

CL jcL =o




В is given by (B.8,3).

(—- ‘j is the slope of the low-speed rolling moment due to yawing at zero ‘ ‘ck, Zо lift, obtained from Fig. B. l 1,1 as a function of aspect ratio, sweep of

the quarter-chord, and taper ratio. (B.11,2) modifies the low-speed value by means of the Prandtl-Glauert rule to yield approximate cor­rections for the first-order three-dimensional effects of compressible flow up to the critical Mach number.

is the increment in Clr due to dihedral, given by
ДС, 1 ttA sin Ac/4

— =——————— — (B.11,3)

Г 12 A + 4 cos A(V4

is the geometric dihedral angle in radians, positive for the wing tip above the plane of the root chord.

is the increment in Clr due to wing twist obtained from Fig. B. l 1,2.

Q is the wing twist between the root and tip sections in degrees, nega­

tive for washout (see Fig. B. l 1,2).

The derivative C„ :


CL is the wing lift coefficient.


~—f is the low-speed drag-due-to-lift yaw-damping parameter obtained from L Fig. B. l 1,3 as a function of wing aspect ratio, taper ratio, sweepback, and CG position.


—— is the low-speed profile-drag yaw-damping parameter obtained from Fig. Do B. l 1,4 as a function of the wing aspect ratio, sweep-back, and CG posi­tion.

CDo is the wing profile drag coefficient evaluated at the appropriate Mach number. For this application Co,, is assumed to be the profile drag associ­ated with the theoretical ideal drag due to lift and is given by

C-Dn Cd

where CD is the total drag coefficient at a given lift coefficient.

Figure B.11,3 Low-speed drag-due-to-lift yaw-damping parameter.