RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 102 AIRFOIL (rae102-il) Reynolds number: 50,000 Max Cl/Cd: 28.4 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae102-il-50000.txt Download as CSV file: xf-rae102-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 102 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6670 0.08480 0.07805 -0.0115 1.0000 0.1739
-8.750 -0.7160 0.07490 0.06814 -0.0189 1.0000 0.1450
-8.500 -0.7450 0.06840 0.06139 -0.0207 1.0000 0.1338
-8.250 -0.7637 0.06240 0.05498 -0.0207 1.0000 0.1253
-8.000 -0.7622 0.05781 0.05003 -0.0198 1.0000 0.1207
-7.750 -0.7577 0.05354 0.04542 -0.0187 1.0000 0.1196
-7.500 -0.7517 0.04955 0.04100 -0.0172 1.0000 0.1196
-7.250 -0.7418 0.04579 0.03677 -0.0155 1.0000 0.1196
-7.000 -0.7281 0.04218 0.03267 -0.0139 1.0000 0.1193
-6.750 -0.7111 0.03885 0.02882 -0.0122 1.0000 0.1202
-6.500 -0.6927 0.03595 0.02541 -0.0106 1.0000 0.1246
-6.250 -0.6703 0.03348 0.02279 -0.0096 1.0000 0.1334
-6.000 -0.6452 0.03082 0.01988 -0.0085 1.0000 0.1408
-5.750 -0.6204 0.02869 0.01772 -0.0075 1.0000 0.1564
-5.500 -0.5935 0.02658 0.01561 -0.0065 1.0000 0.1771
-5.250 -0.5709 0.02451 0.01386 -0.0051 1.0000 0.2131
-5.000 -0.5551 0.02197 0.01200 -0.0028 1.0000 0.2855
-4.750 -0.5614 0.01976 0.01193 0.0052 1.0000 0.5403
-4.500 -0.5575 0.02071 0.01315 0.0144 1.0000 0.6679
-4.250 -0.5410 0.02192 0.01430 0.0217 1.0000 0.7335
-4.000 -0.5077 0.02353 0.01565 0.0269 1.0000 0.7893
-3.750 -0.4178 0.02592 0.01734 0.0235 1.0000 0.8493
-3.500 -0.3194 0.02640 0.01712 0.0132 1.0000 0.8891
-3.250 -0.2614 0.02583 0.01620 0.0075 1.0000 0.9154
-3.000 -0.2068 0.02502 0.01512 0.0016 1.0000 0.9372
-2.750 -0.1587 0.02414 0.01402 -0.0035 1.0000 0.9566
-2.500 -0.1058 0.02308 0.01278 -0.0098 1.0000 0.9734
-2.250 -0.0541 0.02198 0.01153 -0.0161 1.0000 0.9895
-2.000 -0.0141 0.02105 0.01052 -0.0205 1.0000 1.0000
-1.750 -0.0011 0.02054 0.01002 -0.0197 1.0000 1.0000
-1.500 0.0110 0.02007 0.00958 -0.0187 1.0000 1.0000
-1.250 0.0214 0.01966 0.00921 -0.0174 1.0000 1.0000
-1.000 0.0294 0.01930 0.00890 -0.0157 1.0000 1.0000
-0.750 0.0336 0.01901 0.00868 -0.0134 1.0000 1.0000
-0.500 0.0322 0.01880 0.00855 -0.0103 1.0000 1.0000
-0.250 0.0212 0.01870 0.00852 -0.0059 1.0000 1.0000
0.000 0.0000 0.01868 0.00853 0.0000 1.0000 1.0000
0.250 -0.0212 0.01870 0.00852 0.0059 1.0000 1.0000
0.500 -0.0322 0.01880 0.00855 0.0103 1.0000 1.0000
0.750 -0.0336 0.01900 0.00867 0.0134 1.0000 1.0000
1.000 -0.0294 0.01930 0.00890 0.0157 1.0000 1.0000
1.250 -0.0214 0.01966 0.00920 0.0174 1.0000 1.0000
1.500 -0.0109 0.02007 0.00958 0.0187 1.0000 1.0000
1.750 0.0011 0.02053 0.01001 0.0197 1.0000 1.0000
2.000 0.0143 0.02104 0.01051 0.0205 1.0000 1.0000
2.250 0.0543 0.02197 0.01152 0.0161 0.9895 1.0000
2.500 0.1056 0.02306 0.01276 0.0098 0.9735 1.0000
2.750 0.1593 0.02414 0.01401 0.0034 0.9565 1.0000
3.000 0.2071 0.02501 0.01511 -0.0017 0.9372 1.0000
3.250 0.2617 0.02582 0.01619 -0.0075 0.9154 1.0000
3.500 0.3197 0.02638 0.01710 -0.0132 0.8891 1.0000
3.750 0.4181 0.02591 0.01732 -0.0235 0.8493 1.0000
4.000 0.5076 0.02352 0.01565 -0.0269 0.7893 1.0000
4.250 0.5408 0.02192 0.01430 -0.0217 0.7337 1.0000
4.500 0.5572 0.02072 0.01316 -0.0144 0.6683 1.0000
4.750 0.5612 0.01976 0.01193 -0.0052 0.5406 1.0000
5.000 0.5550 0.02197 0.01199 0.0028 0.2854 1.0000
5.250 0.5709 0.02450 0.01386 0.0051 0.2134 1.0000
5.500 0.5934 0.02658 0.01560 0.0065 0.1772 1.0000
5.750 0.6203 0.02869 0.01773 0.0075 0.1564 1.0000
6.000 0.6451 0.03082 0.01988 0.0085 0.1407 1.0000
6.250 0.6702 0.03348 0.02279 0.0096 0.1334 1.0000
6.500 0.6926 0.03594 0.02540 0.0106 0.1247 1.0000
6.750 0.7111 0.03886 0.02883 0.0122 0.1201 1.0000
7.000 0.7282 0.04218 0.03266 0.0139 0.1193 1.0000
7.250 0.7419 0.04580 0.03678 0.0155 0.1196 1.0000
7.500 0.7517 0.04957 0.04103 0.0171 0.1196 1.0000
7.750 0.7579 0.05354 0.04543 0.0186 0.1196 1.0000
8.000 0.7629 0.05782 0.05003 0.0198 0.1209 1.0000
8.250 0.7633 0.06247 0.05506 0.0207 0.1254 1.0000
8.500 0.7454 0.06844 0.06143 0.0206 0.1338 1.0000
8.750 0.7167 0.07501 0.06825 0.0187 0.1455 1.0000
9.000 0.6874 0.08299 0.07629 0.0145 0.1678 1.0000
9.250 0.5431 0.08596 0.07940 0.0080 0.1868 1.0000
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Polar data table (+)
Polar graphs
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