Check your units.
As Man.Utd said, for the conditions you cite the overall CL is 0.27
m - kg
g - m/sec squared
rho - kg/ m cubed
TAS - m/ sec
Area - m squared
Depending on CG the tail lift will be in the range 2 to say 8% of the total so the wing lift will be that much higher. To get an exact value you need detailed aerodynamic information that is almost never published.
The aoa will depend on flap setting and Mach number, but again needs unpublished detail. As a ballpark figure use aoa = 10 times CL (aoa in degrees)