RAE 100 AIRFOIL (rae100-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 100 AIRFOIL (rae100-il) Reynolds number: 100,000 Max Cl/Cd: 35.27 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae100-il-100000.txt Download as CSV file: xf-rae100-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 100 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.8252 0.07540 0.07005 -0.0248 1.0000 0.0730
-10.500 -0.8306 0.07025 0.06484 -0.0254 1.0000 0.0720
-10.250 -0.8487 0.06565 0.06014 -0.0251 1.0000 0.0715
-10.000 -0.8651 0.06098 0.05530 -0.0245 1.0000 0.0709
-9.750 -0.8778 0.05632 0.05037 -0.0236 1.0000 0.0706
-9.500 -0.8865 0.05184 0.04550 -0.0223 1.0000 0.0708
-9.250 -0.8899 0.04769 0.04086 -0.0207 1.0000 0.0715
-9.000 -0.8875 0.04396 0.03659 -0.0188 1.0000 0.0722
-8.750 -0.8799 0.04063 0.03271 -0.0170 1.0000 0.0729
-8.500 -0.8660 0.03669 0.02858 -0.0161 1.0000 0.0748
-8.250 -0.8467 0.03486 0.02672 -0.0153 1.0000 0.0786
-8.000 -0.8295 0.03272 0.02417 -0.0139 1.0000 0.0826
-7.750 -0.8114 0.03048 0.02141 -0.0123 1.0000 0.0857
-7.500 -0.7896 0.02844 0.01947 -0.0117 1.0000 0.0912
-7.250 -0.7682 0.02701 0.01770 -0.0105 1.0000 0.0978
-7.000 -0.7455 0.02517 0.01596 -0.0097 1.0000 0.1047
-6.750 -0.7232 0.02377 0.01438 -0.0087 1.0000 0.1143
-6.500 -0.7007 0.02259 0.01315 -0.0076 1.0000 0.1254
-6.250 -0.6786 0.02138 0.01199 -0.0066 1.0000 0.1397
-6.000 -0.6572 0.02015 0.01090 -0.0055 1.0000 0.1570
-5.750 -0.6363 0.01912 0.01001 -0.0043 1.0000 0.1794
-5.500 -0.6159 0.01813 0.00920 -0.0030 1.0000 0.2084
-5.250 -0.5959 0.01723 0.00848 -0.0016 1.0000 0.2450
-5.000 -0.5762 0.01644 0.00793 -0.0002 1.0000 0.2869
-4.750 -0.5565 0.01577 0.00748 0.0013 1.0000 0.3316
-4.500 -0.5365 0.01522 0.00713 0.0029 1.0000 0.3764
-4.250 -0.5164 0.01476 0.00686 0.0045 1.0000 0.4203
-4.000 -0.4961 0.01437 0.00662 0.0062 1.0000 0.4626
-3.750 -0.4756 0.01404 0.00641 0.0078 1.0000 0.5044
-3.500 -0.4551 0.01373 0.00628 0.0096 1.0000 0.5425
-3.250 -0.4348 0.01347 0.00613 0.0114 1.0000 0.5816
-3.000 -0.4146 0.01322 0.00603 0.0134 1.0000 0.6177
-2.750 -0.3948 0.01300 0.00595 0.0154 1.0000 0.6540
-2.500 -0.3754 0.01281 0.00587 0.0175 1.0000 0.6906
-2.250 -0.3564 0.01264 0.00585 0.0198 1.0000 0.7254
-2.000 -0.3378 0.01251 0.00584 0.0222 1.0000 0.7610
-1.750 -0.3195 0.01242 0.00586 0.0246 1.0000 0.7984
-1.500 -0.2991 0.01241 0.00595 0.0268 1.0000 0.8363
-1.250 -0.2736 0.01250 0.00611 0.0279 1.0000 0.8768
-1.000 -0.2321 0.01273 0.00636 0.0259 1.0000 0.9151
-0.750 -0.1726 0.01302 0.00661 0.0201 1.0000 0.9459
-0.500 -0.1033 0.01323 0.00676 0.0121 1.0000 0.9689
-0.250 -0.0301 0.01326 0.00675 0.0030 1.0000 0.9876
0.000 0.0000 0.01318 0.00667 0.0000 1.0000 1.0000
0.250 0.0301 0.01326 0.00675 -0.0030 0.9876 1.0000
0.500 0.1033 0.01323 0.00676 -0.0121 0.9689 1.0000
0.750 0.1725 0.01302 0.00661 -0.0201 0.9460 1.0000
1.000 0.2321 0.01273 0.00636 -0.0259 0.9151 1.0000
1.250 0.2737 0.01250 0.00611 -0.0279 0.8769 1.0000
1.500 0.2992 0.01241 0.00595 -0.0268 0.8363 1.0000
1.750 0.3196 0.01242 0.00586 -0.0246 0.7983 1.0000
2.000 0.3379 0.01251 0.00584 -0.0222 0.7611 1.0000
2.250 0.3565 0.01264 0.00585 -0.0198 0.7255 1.0000
2.500 0.3756 0.01281 0.00587 -0.0175 0.6906 1.0000
2.750 0.3949 0.01300 0.00595 -0.0154 0.6539 1.0000
3.000 0.4148 0.01322 0.00603 -0.0134 0.6177 1.0000
3.250 0.4349 0.01347 0.00613 -0.0115 0.5816 1.0000
3.500 0.4553 0.01373 0.00628 -0.0097 0.5424 1.0000
3.750 0.4757 0.01404 0.00641 -0.0079 0.5044 1.0000
4.000 0.4962 0.01437 0.00662 -0.0062 0.4626 1.0000
4.250 0.5166 0.01476 0.00686 -0.0046 0.4202 1.0000
4.500 0.5367 0.01522 0.00714 -0.0029 0.3764 1.0000
4.750 0.5566 0.01578 0.00748 -0.0014 0.3315 1.0000
5.000 0.5763 0.01644 0.00793 0.0002 0.2868 1.0000
5.250 0.5960 0.01723 0.00848 0.0016 0.2450 1.0000
5.500 0.6160 0.01813 0.00920 0.0030 0.2083 1.0000
5.750 0.6364 0.01912 0.01001 0.0043 0.1794 1.0000
6.000 0.6572 0.02015 0.01090 0.0055 0.1569 1.0000
6.250 0.6787 0.02138 0.01199 0.0066 0.1397 1.0000
6.500 0.7007 0.02259 0.01314 0.0076 0.1255 1.0000
6.750 0.7232 0.02376 0.01438 0.0087 0.1142 1.0000
7.000 0.7454 0.02516 0.01595 0.0097 0.1047 1.0000
7.250 0.7682 0.02700 0.01769 0.0105 0.0978 1.0000
7.500 0.7895 0.02845 0.01948 0.0117 0.0912 1.0000
7.750 0.8113 0.03049 0.02142 0.0123 0.0857 1.0000
8.000 0.8294 0.03272 0.02418 0.0139 0.0826 1.0000
8.250 0.8465 0.03489 0.02675 0.0153 0.0787 1.0000
8.500 0.8657 0.03672 0.02863 0.0162 0.0749 1.0000
8.750 0.8800 0.04048 0.03256 0.0171 0.0730 1.0000
9.000 0.8874 0.04395 0.03657 0.0189 0.0722 1.0000
9.250 0.8902 0.04767 0.04082 0.0207 0.0716 1.0000
9.500 0.8865 0.05180 0.04545 0.0223 0.0708 1.0000
9.750 0.8786 0.05629 0.05032 0.0236 0.0708 1.0000
10.000 0.8658 0.06095 0.05525 0.0245 0.0710 1.0000
10.250 0.8477 0.06563 0.06013 0.0252 0.0714 1.0000
10.500 0.8341 0.07016 0.06474 0.0255 0.0722 1.0000
10.750 0.8251 0.07524 0.06989 0.0248 0.0731 1.0000
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Polar data table (+)
Polar graphs
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