Center of Pressure Location for Cambered Airfoil Solution

STEP 0: Pre-Calculation Summary
Formula Used
Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient
xcp = -(Cm,le*c)/CL
This formula uses 4 Variables
Variables Used
Center of Pressure - (Measured in Meter) - The Center of Pressure is the point where the total sum of a pressure field acts on a body, causing a force to act through that point.
Moment Coefficient about Leading Edge - Moment Coefficient about Leading Edge is obtained by dividing the moment about the leading edge by the dynamic pressure, the area, and the chord of the airfoil.
Chord - (Measured in Meter) - A chord is the precise distance between the leading and trailing edge of the airfoil.
Lift Coefficient - The Lift Coefficient is a dimensionless coefficient that relates the lift generated by a lifting body to the fluid density around the body, the fluid velocity and an associated reference area.
STEP 1: Convert Input(s) to Base Unit
Moment Coefficient about Leading Edge: -0.3 --> No Conversion Required
Chord: 3 Meter --> 3 Meter No Conversion Required
Lift Coefficient: 1.2 --> No Conversion Required
STEP 2: Evaluate Formula
Substituting Input Values in Formula
xcp = -(Cm,le*c)/CL --> -((-0.3)*3)/1.2
Evaluating ... ...
xcp = 0.75
STEP 3: Convert Result to Output's Unit
0.75 Meter --> No Conversion Required
FINAL ANSWER
0.75 Meter <-- Center of Pressure
(Calculation completed in 00.004 seconds)

Credits

Created by Shikha Maurya
Indian Institute of Technology (IIT), Bombay
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Amrita School of Engineering (ASE), Vallikavu
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8 Flow over Airfoils Calculators

Boundary Layer Thickness for Laminar Flow
Go Laminar Boundary Layer Thickness = 5*Distance on X-Axis/sqrt(Reynolds Number for Laminar Flow)
Center of Pressure Location for Cambered Airfoil
Go Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient
Lift Coefficient for Cambered Airfoil
Go Lift Coefficient for Cambered Airfoil = 2*pi*(Angle of Attack-Angle of Zero Lift)
Boundary Layer Thickness for Turbulent Flow
Go Turbulent Boundary Layer Thickness = 0.37*Distance on X-Axis/(Reynolds Number for Turbulent Flow^(1/5))
Skin Friction Drag Coefficient for Flat Plate in Laminar Flow
Go Skin Friction Drag Coefficient = 1.328/(sqrt(Reynolds Number for Laminar Flow))
Skin Friction Drag Coefficient for Flat Plate in Turbulent Flow
Go Skin Friction Drag Coefficient = 0.074/(Reynolds Number for Turbulent Flow^(1/5))
Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory
Go Lift Coefficient = 2*pi*Angle of Attack
Moment Coefficient about Leading-Edge for Symmetrical Airfoil by Thin Airfoil Theory
Go Moment Coefficient about Leading Edge = -Lift Coefficient/4

Center of Pressure Location for Cambered Airfoil Formula

Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient
xcp = -(Cm,le*c)/CL

What is center of pressure?

Center of pressure is the location where the resultant of a distributed load effectively acts on the body. If moments were taken about the center of pressure, the integrated effect of the distributed loads would be zero. In a cambered airfoil, the aerodynamic center and center of pressure are not at the same point.

How to Calculate Center of Pressure Location for Cambered Airfoil?

Center of Pressure Location for Cambered Airfoil calculator uses Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient to calculate the Center of Pressure, The Center of pressure location for cambered airfoil formula is obtained as the function of moment coefficient about leading edge, lift coefficient and chord of the airfoil. Center of Pressure is denoted by xcp symbol.

How to calculate Center of Pressure Location for Cambered Airfoil using this online calculator? To use this online calculator for Center of Pressure Location for Cambered Airfoil, enter Moment Coefficient about Leading Edge (Cm,le), Chord (c) & Lift Coefficient (CL) and hit the calculate button. Here is how the Center of Pressure Location for Cambered Airfoil calculation can be explained with given input values -> 0.75 = -((-0.3)*3)/1.2.

FAQ

What is Center of Pressure Location for Cambered Airfoil?
The Center of pressure location for cambered airfoil formula is obtained as the function of moment coefficient about leading edge, lift coefficient and chord of the airfoil and is represented as xcp = -(Cm,le*c)/CL or Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient. Moment Coefficient about Leading Edge is obtained by dividing the moment about the leading edge by the dynamic pressure, the area, and the chord of the airfoil, A chord is the precise distance between the leading and trailing edge of the airfoil & The Lift Coefficient is a dimensionless coefficient that relates the lift generated by a lifting body to the fluid density around the body, the fluid velocity and an associated reference area.
How to calculate Center of Pressure Location for Cambered Airfoil?
The Center of pressure location for cambered airfoil formula is obtained as the function of moment coefficient about leading edge, lift coefficient and chord of the airfoil is calculated using Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient. To calculate Center of Pressure Location for Cambered Airfoil, you need Moment Coefficient about Leading Edge (Cm,le), Chord (c) & Lift Coefficient (CL). With our tool, you need to enter the respective value for Moment Coefficient about Leading Edge, Chord & Lift Coefficient and hit the calculate button. You can also select the units (if any) for Input(s) and the Output as well.
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