NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il)
NASA/AMES 63A108 MOD C AIRFOIL - NASA/AMES modification of the NACA 63A-108 airfoil
| Details | Dat file | Parser | |
|
(a63a108c-il) NASA/AMES 63A108 MOD C AIRFOIL NASA/AMES modification of the NACA 63A-108 airfoil Max thickness 7.7% at 30% chord. Max camber 0.5% at 12.5% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
NASA/AMES 63A108 MOD C AIRFOIL 50. 50. .0 .0 .0002 .00445 .0004 .00549 .0006 .00626 .0008 .00687 .001 .00741 .002 .00943 .003 .01091 .004 .01212 .005 .01315 .010 .01695 .02 .02192 .03 .02539 .04 .02807 .05 .03024 .06 .03206 .07 .03360 .08 .03493 .09 .03608 .10 .03709 .125 .03911 .15 .04062 .175 .04173 .20 .04255 .25 .04351 .30 .04385 .35 .04362 .40 .04293 .45 .04176 .50 .04015 .55 .03810 .60 .03559 .65 .03258 .70 .02906 .725 .0272 .75 .0252 .775 .0231 .80 .0209 .825 .0187 .85 .0164 .875 .0141 .90 .0120 .925 .0101 .95 .0086 .96 .0080 .97 .0076 .98 .0073 .99 .0071 .995 .0070 1. .0070 .0 .0 .0002 -.00277 .0004 -.00379 .0006 -.00440 .0008 -.00489 .001 -.00531 .002 -.00686 .003 -.00796 .004 -.00885 .005 -.00959 .010 -.01228 .02 -.01569 .03 -.01808 .04 -.01996 .05 -.02152 .06 -.02286 .07 -.02403 .08 -.02506 .09 -.02598 .10 -.02680 .125 -.02852 .15 -.02989 .175 -.03097 .20 -.03183 .25 -.03297 .30 -.03350 .35 -.03340 .40 -.03269 .45 -.03140 .50 -.02957 .55 -.02728 .60 -.02459 .65 -.02157 .70 -.01830 .725 -.01661 .75 -.01492 .775 -.01327 .80 -.01167 .825 -.01018 .85 -.00874 .875 -.00729 .90 -.00584 .925 -.00438 .95 -.00292 .96 -.00234 .97 -.00176 .98 -.00117 .99 -.00059 .995 -.00029 1. .0 |
No parser warnings |
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| a63a108c-il | 50,000 | 9 | 24.5 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| a63a108c-il | 50,000 | 5 | 22.8 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| a63a108c-il | 100,000 | 9 | 31.7 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| a63a108c-il | 100,000 | 5 | 31.2 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| a63a108c-il | 200,000 | 9 | 37.1 at α=5.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| a63a108c-il | 200,000 | 5 | 41.8 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| a63a108c-il | 500,000 | 9 | 54.4 at α=6.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| a63a108c-il | 500,000 | 5 | 58.8 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| a63a108c-il | 1,000,000 | 9 | 69.5 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| a63a108c-il | 1,000,000 | 5 | 72.7 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||