Calculate Cp Pressure Coefficient at a Point on Wing and Airfoil

 Method to Calculate the Pressure Coefficient Cpat a Point on the Wing and Airfoil

• Calculate Lift about the quarter chord point per unit span

• Calculate Drag about the quarter chord point per unit span

• Calculate Moment about the quarter chord point per unit span

• Calculate Pressure Coefficient at a Point on Wing and Airfoil

• Calculate Induced Drag Coefficient and Induced Drag

• Calculate Total Drag

• Calculate Cl and Cd

• Calculate Lift Coefficient for Airplane

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Pressure Coefficient Cp at a Point on Wing and Airfoil

To calculate the Pressure Coefficient Cp value at a point on the wing we must know the value of the altitude at which the airplane is flying in meters.

We also need to know the pressure at a point on the wing of the airplane in N/m^2

And the value of Airplane Velocity in m/s.

Performing the Calculation we have the following parameters,

The value of Density at given altitude.

The value of Dynamic pressure q = 1/2 * density * velocity^2

and the Value of the Pressure Coefficient Cp

Cp = (p - p(infinity))/ q (infinity)

In order to calculate the pressure Coefficient Cp value at a point on the airfoil placed in a low-speed wind tunnel at standard sea-level conditions we must have,

• Flow Velocity in test section in feet/sec (ft/s)

• Pressure at a point on the airfoil (lb/ft^2)

• and if the flow velocity is increased such that the free stream Mach number becomes value.

The formula used for Pressure coefficient at a point on the wing is Cp = Cp0/ squareroot (1-Mach(infinity)^2)

Download Free Program to Calculate the Pressure Coefficient Cp at a point on an Airfoil or wing

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