Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory Calculator

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Two-Dimensional Incompressible Flow

Compressible Flow

Fundamentals of Inviscid and Incompressible Flow

Three-Dimensional Incompressible Flow

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Flow over Airfoils and Wings

Elementary Flows

Flow and Lift Distribution

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✖Angle of Attack is the angle between a reference line on a body and the vector representing the relative motion between the body and the fluid through which it is moving.ⓘ Angle of Attack [α]

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✖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.ⓘ Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory [C_L]

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Formula

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Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory

Formula

`"C"_{"L"} = 2*pi*"α"`

Example

`"1.199705"=2*pi*"10.94°"`

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Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory Solution

STEP 0: Pre-Calculation Summary

Formula Used

Lift Coefficient = 2*pi*Angle of Attack
C_L = 2*pi*α
This formula uses 1 Constants, 2 Variables

Constants Used

pi - Archimedes' constant Value Taken As 3.14159265358979323846264338327950288

Variables Used

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.
Angle of Attack - (Measured in Radian) - Angle of Attack is the angle between a reference line on a body and the vector representing the relative motion between the body and the fluid through which it is moving.

STEP 1: Convert Input(s) to Base Unit

Angle of Attack: 10.94 Degree --> 0.190939020168144 Radian (Check conversion here)

STEP 2: Evaluate Formula

Substituting Input Values in Formula

C_L = 2*pi*α --> 2*pi*0.190939020168144

Evaluating ... ...

C_L = 1.19970524608775

STEP 3: Convert Result to Output's Unit

1.19970524608775 --> No Conversion Required

FINAL ANSWER

1.19970524608775 ≈ 1.199705 <-- Lift Coefficient

(Calculation completed in 00.006 seconds)

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Home » Physics » Aerodynamics » Two-Dimensional Incompressible Flow » Flow over Airfoils and Wings » Flow over Airfoils » Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory

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Created by Shikha Maurya

Indian Institute of Technology (IIT), Bombay

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Indian Institute for Aeronautical Engineering and Information Technology (IIAEIT), Pune

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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)

Lift Coefficient for Cambered Airfoil

Go Lift Coefficient for Cambered Airfoil = 2*pi*((Angle of Attack)-(Angle of Zero Lift))

Center of Pressure Location for Cambered Airfoil

Go Center of Pressure = -(Moment Coefficient about Leading Edge*Chord)/Lift Coefficient

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

Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory Formula

Lift Coefficient = 2*pi*Angle of Attack
C_L = 2*pi*α

What is thin airfoil theory?

Thin airfoil theory is predicated on the replacement of the airfoil by the mean camber line. A vortex sheet is placed along the chord line and its strength adjusted such that, in conjunction with the uniform freestream, the camber line becomes a streamline of the flow while at the same time satisfying the Kutta condition.

What is Kutta Condition?

The Kutta condition is an observation that for a lifting airfoil of a given shape at a
given angle of attack, nature adopts that particular value of circulation around the airfoil which results in the flow leaving smoothly at the trailing edge. If the trailing-edge angle is finite, then the trailing edge is a stagnation point. If the trailing edge is cusped, then the velocities leaving the top and bottom surfaces at the trailing edge are finite and equal in magnitude and direction.

How to Calculate Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory?

Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory calculator uses Lift Coefficient = 2*pi*Angle of Attack to calculate the Lift Coefficient, The Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory formula, the lift coefficient for a symmetrical airfoil is determined by the angle of attack, The lift coefficient increases linearly with the angle of attack according to the formula. Lift Coefficient is denoted by C_L symbol.

How to calculate Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory using this online calculator? To use this online calculator for Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory, enter Angle of Attack (α) and hit the calculate button. Here is how the Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory calculation can be explained with given input values -> 1.199705 = 2*pi*0.190939020168144.

FAQ

What is Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory?

The Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory formula, the lift coefficient for a symmetrical airfoil is determined by the angle of attack, The lift coefficient increases linearly with the angle of attack according to the formula and is represented as C_L = 2*pi*α or Lift Coefficient = 2*pi*Angle of Attack. Angle of Attack is the angle between a reference line on a body and the vector representing the relative motion between the body and the fluid through which it is moving.

How to calculate Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory?

The Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory formula, the lift coefficient for a symmetrical airfoil is determined by the angle of attack, The lift coefficient increases linearly with the angle of attack according to the formula is calculated using Lift Coefficient = 2*pi*Angle of Attack. To calculate Lift Coefficient for Symmetrical Airfoil by Thin Airfoil Theory, you need Angle of Attack (α). With our tool, you need to enter the respective value for Angle of Attack 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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