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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.

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 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.

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:

$$ 2400 = \frac {1}{2} \cdot \rho \cdot \Delta V^2 => \Delta V = \sqrt{\frac{2400 \cdot 2 }{ 1.225}} = 62.6 m/s$$

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 162.6 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.

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.

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. 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.

Image source

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:

$$ 2400 = \frac {1}{2} \cdot \rho \cdot \Delta V^2 => \Delta V = \sqrt{\frac{2400 \cdot 2 }{ 1.225}} = 62.6 m/s$$

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 162.6 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.