If CM0 is -0.052 and lift coefficient at zero angle CL0 is 0.92 then, find CMac. Consider rectangular wing.
(a) -0.282
(b) 0.98
(c) 2.5
(d) 7.89
The question was posed to me during an online interview.
My doubt stems from Longitudinal Static Stability and Control-2 topic in portion Stability, Control, and Handling Qualities of Aircraft Design