## Drag on Airfoil Solution

STEP 0: Pre-Calculation Summary
Formula Used
Drag on Airfoil = Normal Force on Airfoil*sin(Angle of Attack of Airfoil)+Axial Force on Airfoil*cos(Angle of Attack of Airfoil)
D = N*sin(α°)+A*cos(α°)
This formula uses 2 Functions, 4 Variables
Functions Used
sin - Sine is a trigonometric function that describes the ratio of the length of the opposite side of a right triangle to the length of the hypotenuse., sin(Angle)
cos - Cosine of an angle is the ratio of the side adjacent to the angle to the hypotenuse of the triangle., cos(Angle)
Variables Used
Drag on Airfoil - (Measured in Newton) - Drag on Airfoil is the component of resultant Force acting on airfoil parallel to freestream velocity.
Normal Force on Airfoil - (Measured in Newton) - Normal Force on Airfoil is component of resultant force acting on airfoil perpendicular to chord.
Angle of Attack of Airfoil - (Measured in Radian) - The Angle of Attack of Airfoil is the angle between freestream velocity and chord of the airfoil.
Axial Force on Airfoil - (Measured in Newton) - Axial Force on Airfoil is component of resultant force acting on airfoil parallel to chord.
STEP 1: Convert Input(s) to Base Unit
Normal Force on Airfoil: 11 Newton --> 11 Newton No Conversion Required
Angle of Attack of Airfoil: 8 Degree --> 0.13962634015952 Radian (Check conversion ​here)
Axial Force on Airfoil: 20 Newton --> 20 Newton No Conversion Required
STEP 2: Evaluate Formula
Substituting Input Values in Formula
D = N*sin(α°)+A*cos(α°) --> 11*sin(0.13962634015952)+20*cos(0.13962634015952)
Evaluating ... ...
D = 21.3362654853919
STEP 3: Convert Result to Output's Unit
21.3362654853919 Newton --> No Conversion Required
21.3362654853919 21.33627 Newton <-- Drag on Airfoil
(Calculation completed in 00.004 seconds)
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## Credits

Created by Vishal Anand
Indian Institute of Technology Kharagpur (IIT KGP), Kharagpur
Vishal Anand has created this Calculator and 7 more calculators!
Verified by Ayush Singh
Gautam Buddha University (GBU), Greater Noida
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## < 7 Computational Fluid Dyanmics Calculators

Drag on Airfoil
Drag on Airfoil = Normal Force on Airfoil*sin(Angle of Attack of Airfoil)+Axial Force on Airfoil*cos(Angle of Attack of Airfoil)
Lift on Airfoil
Lift on Airfoil = Normal Force on Airfoil*cos(Angle of Attack of Airfoil)-Axial Force on Airfoil*sin(Angle of Attack of Airfoil)
Reynolds Number for Airfoil
Reynolds Number = (Density of Fluid*Flow Velocity*Chord Length of Airfoil)/Dynamic Viscosity
Wall Shear Stress for Airfoil
Wall Shear Stress For Airfoil = 0.5*Skin Friction Coefficient*Flow Velocity^2*Density of Air
Y Plus
Y Plus = (First Layer Height*Friction Velocity For Airfoil)/Kinematic Viscosity
Friction Velocity for Airfoil
Friction Velocity For Airfoil = (Wall Shear Stress For Airfoil/Density of Air)^0.5
Skin Friction Coefficient
Skin Friction Coefficient = (2*log10(Reynolds Number)-0.65)^(-2.30)

## Drag on Airfoil Formula

Drag on Airfoil = Normal Force on Airfoil*sin(Angle of Attack of Airfoil)+Axial Force on Airfoil*cos(Angle of Attack of Airfoil)
D = N*sin(α°)+A*cos(α°)

## How angle of attack effects drag?

The magnitude of the drag generated by an object depends on the shape of the object and how it moves through the air. For airfoils, the drag is nearly constant at small angles (+/- 5 degrees). As the angle increases above 5 degrees, the drag quickly rises because of increased frontal area and increased boundary layer thickness.

## How to Calculate Drag on Airfoil?

Drag on Airfoil calculator uses Drag on Airfoil = Normal Force on Airfoil*sin(Angle of Attack of Airfoil)+Axial Force on Airfoil*cos(Angle of Attack of Airfoil) to calculate the Drag on Airfoil, The Drag on Airfoil is the aerodynamic force that opposes the motion of the airfoil through the air. It's caused by the friction and pressure differences between the upper and lower surfaces of the airfoil as air flows around it. Drag is a critical consideration in aircraft design as it affects fuel efficiency and performance. Drag on Airfoil is denoted by D symbol.

How to calculate Drag on Airfoil using this online calculator? To use this online calculator for Drag on Airfoil, enter Normal Force on Airfoil (N), Angle of Attack of Airfoil °) & Axial Force on Airfoil (A) and hit the calculate button. Here is how the Drag on Airfoil calculation can be explained with given input values -> 21.33627 = 11*sin(0.13962634015952)+20*cos(0.13962634015952).

### FAQ

What is Drag on Airfoil?
The Drag on Airfoil is the aerodynamic force that opposes the motion of the airfoil through the air. It's caused by the friction and pressure differences between the upper and lower surfaces of the airfoil as air flows around it. Drag is a critical consideration in aircraft design as it affects fuel efficiency and performance and is represented as D = N*sin(α°)+A*cos(α°) or Drag on Airfoil = Normal Force on Airfoil*sin(Angle of Attack of Airfoil)+Axial Force on Airfoil*cos(Angle of Attack of Airfoil). Normal Force on Airfoil is component of resultant force acting on airfoil perpendicular to chord, The Angle of Attack of Airfoil is the angle between freestream velocity and chord of the airfoil & Axial Force on Airfoil is component of resultant force acting on airfoil parallel to chord.
How to calculate Drag on Airfoil?
The Drag on Airfoil is the aerodynamic force that opposes the motion of the airfoil through the air. It's caused by the friction and pressure differences between the upper and lower surfaces of the airfoil as air flows around it. Drag is a critical consideration in aircraft design as it affects fuel efficiency and performance is calculated using Drag on Airfoil = Normal Force on Airfoil*sin(Angle of Attack of Airfoil)+Axial Force on Airfoil*cos(Angle of Attack of Airfoil). To calculate Drag on Airfoil, you need Normal Force on Airfoil (N), Angle of Attack of Airfoil °) & Axial Force on Airfoil (A). With our tool, you need to enter the respective value for Normal Force on Airfoil, Angle of Attack of Airfoil & Axial Force on Airfoil 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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