Calculate CL and CL of Wing

 Method to Calculate the Lift Coefficient CD and Drag Coefficient CD of Wing

• 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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Lift and Drag Coefficient CL and CD of Wing

In order to calculate the Lift Coefficient Cl and Drag Coefficient Cd of an Wing we must have the Aspect Ratio of the Wing.

The Reynolds Number of the flow in kg / m*s units.

The Span Efficiency Factor, e Value

The Angle of Attack alpha (AOA) value at which the lift and drag coefficient are to be calculated.

At effective angle of attack of zero and ten degrees the value of the corresponding lift coefficient.

The angle of attack value at zero lift and the profile drag coefficient Cd value.

Performing the calculation, we get the following parameters,

The lift curve slope value in per degree units

At the given Angle of Attack the CL value of Wing.

At the given Angle of Attack the CD value of Wing.

Download Free Program to Calculate the Lift CL coefficient and Drag CD coefficient of a Wing