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W=2.27*10^4;//weight of the airplane(N)
S=19;//wing area (m^2)
V=61;//velocity at sea level(m/s)
D=1.225;//density at sea level(Kg/m^3)
Cl=2*W/(D*S*V^2) //lift coefficient
a=0.08;//lift slope per degree (from example 7.3)
a1=Cl/a //absolute angle of attack
DCmcg=-0.0133;//derivative of Cmcg w.r.t absolute angle of attack(from example 7.5)
Cmo=0.06;//value of moment coefficient at zero absolute angle of attack (from example 7.5)
Vh=0.34 //tail volume ratio(from example 7.4)
DClt=0.04;//elevator control efficiency
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