Similarly shaped airfoils to NACA67-411 a=0.2
New! Bulk download of polars in Excel format: NACA67-411%20a%3D0.2.xlsx
Info | DAT and DXF files | Shape | Venkataraman Fit | Applications (0) | Similarly shaped airfoils
Plain Flapped Shapes: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Basic Data: cl vs α | cd vs α | cm vs α | cl/cd vs α | cd vs cl | cl/cd vs cl | cl vs cd | cm vs cl | Curve fits | Lift curve slopes
Data source comparison (Codes, CFD, Wind tunnel)
Compressibility Effects: cl vs α | cd vs α | cm vs α | cl/cd vs α | cd vs cl | cl/cd vs cl | cl vs cd | cm vs cl
Roughness Effects: cl vs α | cd vs α | cm vs α | cl/cd vs α | cd vs cl | cl/cd vs cl | cl vs cd | cm vs cl
Plain Flap Effects: cl,max vs flap deflection: all flap chords | 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Plain Flap Effects: cl vs α for various flap deflections: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Plain Flap Effects: cm vs α for various flap deflections: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Plain Flap Effects: cd vs α for various flap deflections: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Disclaimer: All information presented herein is delivered without guarantee or warranty of any kind. The user assumes the entire risk of use of this information.
2. NACA67-511 a=0.0
3. NACA66-411 a=0.2
4. NACA66-411 a=0.0
5. NACA67-311 a=0.5
6. NACA67-411 a=0.5
7. NACA67-412 a=0.2
8. NACA66-412 a=0.2
9. NACA67-410 a=0.2
10. NACA67-410 a=0.0
11. NACA67-311 a=0.6
12. NACA66-412 a=0.0
13. NACA67-412 a=0.0
14. NACA66-511 a=0.0
15. NACA67-311 a=0.2
16. NACA67-511 a=0.2
17. NACA67-512 a=0.0
18. NACA0012-64 a=0.8 c(li)=0.2
19. NACA66-512 a=0.0
20. NACA67-411 a=0.6
21. NACA66-311 a=0.5