OK if it is lift that you would like to compute, this is the way it is commonly done:
$$ L = C_L * \frac {1}{2} \cdot \rho \cdot V^2 \cdot A $$
With
- L= lift [N]
- $C_L$ a lift coefficient depending on aerofoil shape and angle of attack $\alpha$ [dimensionless]
- $\rho$ = air density [kg/$m^3$]
- V = airspeed (of the aircraft) [m/s]
- A = wing area [$m^2$]
If you would like to use a flat bottomed profile, you could use Clark Y. This image is a bit fuzzy but you could construct a straight line $C_L - \alpha$ from the graph.
Image source
At $\alpha$ = 0 the $C_L$ is about 0.4, so at sea level, with an airspeed of 100 m/s and a wing area of 10 $m^2$ the lift would be:
$$ L = 0.4 \cdot \frac {1}{2} * 1.225 * 100^2 * 10 $$
= 24,500 N. So the pressure differential over the wing area is 2,450 N/m$^2$ If this would be from increased dynamic pressure over the wing only, the $\Delta V$ on top of the wing would be:
$$ 2,450 = \frac {1}{2} \cdot \rho \cdot \Delta V^2 => \Delta V = \sqrt{\frac{2,450 \cdot 2 }{ 1.225}} = 63.2 m/s$$
In that particular case, at angle of attack zero, with no skin friction or boundary layer etc etc, the speed under the flat wing surface would be the airspeed = 100 m/s and over the wing would be 163.2 m/s. But the creation of aerodynamic lift is way more complicated than only considering the speeds over and under a wing section and assuming that lower surface pressure = atmospheric pressure, that is why it is done with the $C_L - C_D$ graphs.