Roughly 1.2 at an angle of 14 degrees (depends on Reynolds and mach numbers). Going over this angle will stall the profile.
Yes
min lift coeff of 1412 =.60 thefore Cl will be 1.10 at angle of attatck 4 and mach .8
For cylinders coefficient of lift is approximately half of coefficient of drag while they are equal for Aerofoils.
coefficient of drag in 0 lift
0.08
0.016
It depends on the Reduced Velocity and amplitude of oscillation. Lift Coefficient could be as high as 1.0, and as low as -10.0 at very low reduced velocities.
0.032
A wing will generate lift according to the following equation: L = ½ A C ρ v² A = wing area C = lift coefficient ρ = air density v = air speed The lift coefficient C is a function of Angle of Attack (AOA), which is the angle between the wing's chord line and the relative wind. The greater the angle, the greater the lift coefficient up until the critical AOA where the wing begins to stall and lose lift. The lift coefficient is also a function of wing aspect ratio and will be specific to a certain airfoil shape.
I'm not sure if I understand you question but Lift Coefficient refers to the lifting force of a wing. Engines do not provide Lift; only Thrust.
For no lift, The induced drag will be zero. However, there will still be drag due to viscous forces and pressure forces.
LIft = coefficient times density times velocity squared times wing area divided by 2 drag= coefficient times density times velocity squared over 2 times reference area