- #1
RandomGuy88
- 404
- 6
Does anyone know how to determine the leading-edge radius of an airfoil using just the x y coordinates.
Can the Leading-Edge Radius of an Airfoil be Determined from X Y Coordinates?
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SUMMARY
The leading-edge radius of an airfoil can be determined from its x-y coordinates by utilizing a curve fitting approach and analytical methods. Specifically, XFoil can be employed to analyze NACA 4-series airfoils, where the leading-edge radius (RLE) is calculated using the formula RLE = 1.1019t2, with 't' representing the maximum thickness derived from the thickness distribution equation. For NACA 0012, users can execute a series of commands in XFoil to obtain necessary parameters, although the method may vary for non-NACA profiles.
PREREQUISITES- Understanding of NACA airfoil series, specifically NACA 4-series
- Familiarity with XFoil software for airfoil analysis
- Knowledge of curve fitting techniques for data analysis
- Basic grasp of aerodynamic principles related to airfoil design
- Learn how to perform curve fitting using Python libraries like NumPy or SciPy
- Explore advanced features of XFoil for analyzing different airfoil profiles
- Research the implications of using non-NACA profiles in aerodynamic calculations
- Study the thickness distribution equations for various airfoil designs
Aerospace engineers, aerodynamicists, and students involved in airfoil design and analysis will benefit from this discussion, particularly those working with XFoil and NACA airfoil profiles.
- #2
- 3,481
- 1,291
You could just do a curve fit and then find it analytically...
- #3
AIR&SPACE
- 100
- 0
I know you can do it in XFoil...
ie:
1) Start XFoil
2) type "NACA 0012"
3) type "OPER"
4) hit "enter/return" key
5) type "GDES"
6) a window pops up:
ie:
1) Start XFoil
2) type "NACA 0012"
3) type "OPER"
4) hit "enter/return" key
5) type "GDES"
6) a window pops up:
- #4
RandomGuy88
- 404
- 6
Do you know if there is a way to output the value of rLE from XFOIL. I need the value of the leading edge radius for a code I am writing and would prefer to not have to enter that value manually..
- #5
AIR&SPACE
- 100
- 0
I knew I'd find it eventually... had to find my book first.
Essentially, if you have a NACA 4 series then NACA ABCD, where CD gives the max thickness, t.
Then: R_{LE}=1.1019t^{2}
Essentially, if you have a NACA 4 series then NACA ABCD, where CD gives the max thickness, t.
Then: R_{LE}=1.1019t^{2}
- #6
AIR&SPACE
- 100
- 0
And if all you have is x-y coordinates then you should be able to find 't' by the thickness distribution (assuming you know it's a NACA 4-series)
\pm y_{t}=\frac{t}{0.20}\left(0.29690 \sqrt{x}-0.12600x-0.35160x^{2}+0.28430x^{3}-0.10150x^{4}\right)
The source: Theory of wing sections: including a summary of airfoil data By Ira H. Abbott, Albert Edward Von Doenhoff
\pm y_{t}=\frac{t}{0.20}\left(0.29690 \sqrt{x}-0.12600x-0.35160x^{2}+0.28430x^{3}-0.10150x^{4}\right)
The source: Theory of wing sections: including a summary of airfoil data By Ira H. Abbott, Albert Edward Von Doenhoff
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- #7
akallabeth
- 1
- 0
Hey! had more or less the same problem, do you think the equation given by Abbott gives an appropriate value if it's not a NACA four digit profile ? For example for normal nowadays aircrafts ? (I don't know what kind of profile I have)
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