RAF 6
New! Bulk download of polars in Excel format: raf6.xlsx
Info | DAT and DXF files | Shape | Venkataraman Fit | Applications (11) | 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.
DXF file of spline and points: download
Lednicer-format DAT file download
VisualFoil-format DAT file download
X,Y,Z CSV file download
Selig-format DAT file (download):
RAF 6 AIRFOIL
1.00000000 0.00150000
0.89988000 0.03314000
0.79981010 0.05378000
0.69974000 0.07142000
0.59970000 0.08406000
0.49967000 0.09170000
0.39966000 0.09534000
0.29966000 0.09578000
0.19967000 0.09062000
0.09973000 0.07426000
0.04981000 0.05408000
0.02487000 0.03599000
0.00000000 0.00000000
0.00000000 0.00000000
0.02502000 -0.00501000
0.05002000 -0.00492000
0.10002000 -0.00474000
0.20002000 -0.00438000
0.30001000 -0.00402000
0.40001000 -0.00366000
0.50001000 -0.00330000
0.60001000 -0.00294000
0.70001000 -0.00258000
0.80001000 -0.00222000
0.90001000 -0.00186000
1.00000000 -0.00150000