Question: Assuming a flat bottom wing moving through the air at say 50km/h and at zero degree angle of attack, what will be the likely airspeed above and below the wing?

I know this will be affected by camber and maybe wing chord too. But is there like a proportional estimate for specific wing camber?

I haven't got much knowledge about wing maths, but a formula would be appreciated.

Applying Bernoulli's equation requires the relative airflow above and below the wing. The things is, I'm trying to experiment with the fact that given atmospheric pressure, the two different airflows, and the wing area, one can actually determine the pressure difference between the top and bottom surface using Bernoulli's equation and also the total force holding the plane up by getting the product of the pressure difference and the wing area. So what I'm trying to do is estimate the lift force on my theoretical airplane model at a particular wind speed. I want to use this to estimate the minimum airspeed required for normal flight 'without flaps'.

Answer: 
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. 
[![enter image description here](https://i.stack.imgur.com/SmDrz.gif)](https://i.stack.imgur.com/SmDrz.gif)[Image source](http://marlongofast.tripod.com/zipped_aeronotes/clarky.htm)

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 air speed on top of the wing would be:

$$ 2,450 = \frac {1}{2} \cdot \rho \cdot (V\_{top}^2 - V\_{bottom}^2) => V\_{top} = \sqrt{\frac{2,450 \cdot 2 }{ 1.225} + 100^2} = 118.3 m/s$$

In that particular case, at angle of attack zero, with no skin friction or boundary layer, static pressure underneath the wing equal to atmospheric pressure 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.
Applying Bernoulli's equation requires the relative airflow above and below the wing. The things is, I'm trying to experiment with the fact that given atmospheric pressure, the two different airflows, and the wing area, one can actually determine the pressure difference between the top and bottom surface using Bernoulli's equation and also the total force holding the plane up by getting the product of the pressure difference and the wing area. 
 So what I'm trying to do is estimate the lift force on my theoretical airplane model at a particular wind speed. I want to use this to estimate the minimum airspeed required for normal flight 'without flaps'.