RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 26 AIRFOIL (raf26-il) Reynolds number: 50,000 Max Cl/Cd: 33.96 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf26-il-50000.txt Download as CSV file: xf-raf26-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 26 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4879 0.09609 0.08928 -0.0087 1.0000 0.2499
-7.500 -0.5066 0.09528 0.08863 -0.0090 1.0000 0.2597
-7.250 -0.5040 0.09204 0.08547 -0.0080 1.0000 0.2735
-6.750 -0.4881 0.08505 0.07858 -0.0040 1.0000 0.3134
-6.500 -0.4785 0.08156 0.07508 -0.0016 1.0000 0.3358
-6.250 -0.4917 0.08007 0.07372 -0.0015 1.0000 0.3571
-6.000 -0.4748 0.07600 0.06968 0.0021 1.0000 0.3823
-5.750 -0.4678 0.07270 0.06643 0.0045 1.0000 0.4049
-5.500 -0.4649 0.06986 0.06367 0.0070 1.0000 0.4310
-4.750 -0.4029 0.04199 0.03381 -0.0414 1.0000 0.1334
-4.500 -0.3796 0.03760 0.02846 -0.0417 1.0000 0.1213
-4.250 -0.3581 0.03411 0.02466 -0.0409 1.0000 0.1189
-4.000 -0.3349 0.03122 0.02119 -0.0400 1.0000 0.1196
-3.750 -0.3128 0.02885 0.01859 -0.0390 1.0000 0.1282
-3.500 -0.2876 0.02662 0.01582 -0.0378 1.0000 0.1336
-3.250 -0.2629 0.02453 0.01354 -0.0367 1.0000 0.1436
-3.000 -0.2386 0.02286 0.01181 -0.0356 1.0000 0.1689
-2.750 -0.2122 0.02111 0.00997 -0.0345 1.0000 0.2053
-2.500 -0.1861 0.01874 0.00853 -0.0337 1.0000 0.3384
-2.250 -0.1343 0.01554 0.00733 -0.0347 1.0000 1.0000
-2.000 -0.1126 0.01554 0.00679 -0.0339 1.0000 1.0000
-1.750 -0.0913 0.01558 0.00640 -0.0330 1.0000 1.0000
-1.500 -0.0701 0.01565 0.00607 -0.0321 1.0000 1.0000
-1.250 -0.0491 0.01575 0.00586 -0.0312 1.0000 1.0000
-1.000 -0.0282 0.01588 0.00572 -0.0302 1.0000 1.0000
-0.750 -0.0075 0.01603 0.00565 -0.0293 1.0000 1.0000
-0.500 0.0131 0.01621 0.00563 -0.0284 1.0000 1.0000
-0.250 0.0335 0.01642 0.00564 -0.0275 1.0000 1.0000
0.000 0.0537 0.01665 0.00573 -0.0266 1.0000 1.0000
0.250 0.0738 0.01692 0.00588 -0.0257 1.0000 1.0000
0.500 0.0936 0.01721 0.00608 -0.0249 1.0000 1.0000
0.750 0.1133 0.01753 0.00633 -0.0241 1.0000 1.0000
1.000 0.1328 0.01789 0.00663 -0.0233 1.0000 1.0000
1.250 0.1520 0.01829 0.00699 -0.0226 1.0000 1.0000
1.500 0.1710 0.01872 0.00740 -0.0218 1.0000 1.0000
1.750 0.1897 0.01918 0.00787 -0.0212 1.0000 1.0000
2.000 0.2082 0.01969 0.00840 -0.0205 1.0000 1.0000
2.250 0.2265 0.02025 0.00902 -0.0200 1.0000 1.0000
2.500 0.2444 0.02085 0.00968 -0.0194 1.0000 1.0000
2.750 0.2619 0.02151 0.01041 -0.0190 1.0000 1.0000
3.000 0.2791 0.02224 0.01122 -0.0186 1.0000 1.0000
3.250 0.2957 0.02303 0.01212 -0.0183 1.0000 1.0000
3.500 0.3117 0.02392 0.01314 -0.0181 1.0000 1.0000
3.750 0.3476 0.02522 0.01472 -0.0220 0.9896 1.0000
4.000 0.4534 0.02651 0.01664 -0.0373 0.9395 1.0000
4.250 0.6042 0.02409 0.01550 -0.0535 0.8613 1.0000
4.500 0.6876 0.02051 0.01290 -0.0541 0.7721 1.0000
4.750 0.6986 0.02057 0.01085 -0.0415 0.3457 1.0000
5.000 0.7100 0.02376 0.01270 -0.0386 0.2209 1.0000
5.250 0.7368 0.02683 0.01520 -0.0379 0.1587 1.0000
5.500 0.7701 0.02965 0.01799 -0.0376 0.1297 1.0000
5.750 0.8000 0.03238 0.02091 -0.0370 0.1149 1.0000
6.000 0.8291 0.03550 0.02434 -0.0361 0.1092 1.0000
6.250 0.8531 0.03897 0.02813 -0.0349 0.1048 1.0000
6.500 0.8718 0.04202 0.03192 -0.0328 0.1011 1.0000
6.750 0.8887 0.04575 0.03620 -0.0308 0.1015 1.0000
7.000 0.9042 0.05008 0.04096 -0.0289 0.1044 1.0000
7.250 0.9122 0.05447 0.04617 -0.0263 0.1117 1.0000
7.500 0.9267 0.06006 0.05193 -0.0252 0.1187 1.0000
7.750 0.9205 0.06533 0.05812 -0.0231 0.1358 1.0000
8.000 0.9245 0.07418 0.06742 -0.0242 0.1737 1.0000
8.250 0.8787 0.08144 0.07507 -0.0282 0.2051 1.0000
8.500 0.7956 0.07554 0.06949 -0.0188 0.1755 1.0000
8.750 0.7558 0.08170 0.07561 -0.0213 0.1776 1.0000
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Polar data table (+)
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