RAF 31 AIRFOIL (raf31-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 31 AIRFOIL (raf31-il) Reynolds number: 50,000 Max Cl/Cd: 33.49 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf31-il-50000-n5.txt Download as CSV file: xf-raf31-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 31 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4886 0.09865 0.09086 -0.0464 1.0000 0.0731
-10.750 -0.5000 0.09240 0.08468 -0.0493 1.0000 0.0727
-10.500 -0.5194 0.08448 0.07683 -0.0535 1.0000 0.0721
-10.250 -0.5497 0.07579 0.06819 -0.0590 1.0000 0.0712
-10.000 -0.5881 0.06867 0.06109 -0.0628 1.0000 0.0704
-9.750 -0.6240 0.06427 0.05667 -0.0620 1.0000 0.0698
-9.500 -0.6505 0.06053 0.05281 -0.0599 1.0000 0.0696
-9.250 -0.6693 0.05713 0.04924 -0.0575 1.0000 0.0697
-9.000 -0.6823 0.05405 0.04597 -0.0549 1.0000 0.0700
-8.750 -0.6904 0.05129 0.04297 -0.0521 1.0000 0.0707
-8.500 -0.6950 0.04879 0.04020 -0.0493 1.0000 0.0719
-8.250 -0.6970 0.04642 0.03750 -0.0464 1.0000 0.0733
-8.000 -0.6962 0.04412 0.03482 -0.0436 1.0000 0.0749
-7.750 -0.6918 0.04199 0.03224 -0.0410 1.0000 0.0761
-7.500 -0.6827 0.04000 0.03006 -0.0389 1.0000 0.0773
-7.250 -0.6712 0.03832 0.02827 -0.0369 1.0000 0.0786
-7.000 -0.6588 0.03682 0.02663 -0.0350 1.0000 0.0803
-6.750 -0.6454 0.03542 0.02507 -0.0331 1.0000 0.0822
-6.500 -0.6311 0.03411 0.02357 -0.0313 1.0000 0.0846
-6.250 -0.6163 0.03295 0.02214 -0.0295 1.0000 0.0884
-6.000 -0.6011 0.03184 0.02093 -0.0278 1.0000 0.0924
-5.750 -0.5863 0.03090 0.01999 -0.0261 1.0000 0.0970
-5.500 -0.5685 0.02994 0.01887 -0.0247 0.9994 0.1027
-5.250 -0.5384 0.02883 0.01775 -0.0258 0.9940 0.1109
-5.000 -0.5097 0.02777 0.01665 -0.0267 0.9881 0.1240
-4.750 -0.4804 0.02684 0.01577 -0.0278 0.9823 0.1454
-4.500 -0.4516 0.02602 0.01508 -0.0290 0.9760 0.1783
-4.250 -0.4232 0.02521 0.01457 -0.0302 0.9701 0.2311
-4.000 -0.3950 0.02460 0.01433 -0.0312 0.9638 0.3030
-3.750 -0.3688 0.02427 0.01421 -0.0314 0.9572 0.3717
-3.500 -0.3409 0.02406 0.01408 -0.0318 0.9508 0.4239
-3.250 -0.3128 0.02393 0.01396 -0.0320 0.9446 0.4657
-3.000 -0.2856 0.02386 0.01394 -0.0320 0.9379 0.5063
-2.750 -0.2565 0.02386 0.01402 -0.0321 0.9322 0.5558
-2.500 -0.2341 0.02387 0.01414 -0.0308 0.9248 0.6025
-2.250 -0.2031 0.02391 0.01424 -0.0310 0.9194 0.6475
-2.000 -0.1798 0.02391 0.01426 -0.0299 0.9114 0.6820
-1.750 -0.1448 0.02388 0.01421 -0.0308 0.9049 0.7176
-1.500 -0.1171 0.02383 0.01415 -0.0304 0.8950 0.7530
-1.250 -0.0743 0.02373 0.01404 -0.0324 0.8884 0.7927
-1.000 -0.0424 0.02369 0.01401 -0.0326 0.8781 0.8317
-0.750 0.0150 0.02365 0.01392 -0.0376 0.8735 0.8695
-0.500 0.0648 0.02379 0.01401 -0.0419 0.8661 0.9036
-0.250 0.1294 0.02386 0.01398 -0.0491 0.8617 0.9333
0.000 0.1960 0.02384 0.01385 -0.0568 0.8582 0.9597
0.250 0.2460 0.02394 0.01388 -0.0619 0.8500 0.9906
0.500 0.2848 0.02388 0.01372 -0.0647 0.8435 1.0000
0.750 0.2954 0.02402 0.01377 -0.0623 0.8313 1.0000
1.000 0.3131 0.02416 0.01383 -0.0610 0.8207 1.0000
1.250 0.3438 0.02416 0.01376 -0.0617 0.8130 1.0000
1.500 0.3589 0.02441 0.01395 -0.0598 0.8011 1.0000
1.750 0.3813 0.02458 0.01408 -0.0589 0.7911 1.0000
2.000 0.4114 0.02462 0.01410 -0.0592 0.7826 1.0000
2.250 0.4300 0.02486 0.01432 -0.0577 0.7706 1.0000
2.500 0.4553 0.02496 0.01442 -0.0571 0.7597 1.0000
2.750 0.4890 0.02484 0.01432 -0.0576 0.7506 1.0000
3.000 0.5091 0.02505 0.01453 -0.0561 0.7377 1.0000
3.250 0.5324 0.02519 0.01469 -0.0551 0.7253 1.0000
3.500 0.5606 0.02519 0.01475 -0.0547 0.7142 1.0000
3.750 0.5915 0.02510 0.01470 -0.0546 0.7034 1.0000
4.250 0.6350 0.02547 0.01520 -0.0519 0.6760 1.0000
4.500 0.6586 0.02565 0.01544 -0.0509 0.6633 1.0000
4.750 0.6869 0.02573 0.01561 -0.0505 0.6525 1.0000
5.000 0.7120 0.02590 0.01589 -0.0497 0.6404 1.0000
5.250 0.7330 0.02616 0.01625 -0.0482 0.6261 1.0000
5.500 0.7575 0.02609 0.01627 -0.0469 0.6094 1.0000
5.750 0.7770 0.02599 0.01624 -0.0446 0.5875 1.0000
6.000 0.7997 0.02558 0.01582 -0.0424 0.5613 1.0000
6.250 0.8154 0.02556 0.01581 -0.0396 0.5327 1.0000
6.500 0.8332 0.02565 0.01590 -0.0373 0.5056 1.0000
6.750 0.8529 0.02578 0.01603 -0.0352 0.4790 1.0000
7.000 0.8714 0.02602 0.01619 -0.0331 0.4502 1.0000
7.250 0.8867 0.02648 0.01656 -0.0307 0.4196 1.0000
7.500 0.8971 0.02713 0.01719 -0.0278 0.3867 1.0000
7.750 0.9030 0.02790 0.01799 -0.0244 0.3486 1.0000
8.000 0.9074 0.02876 0.01872 -0.0209 0.3072 1.0000
8.250 0.9106 0.02988 0.01949 -0.0174 0.2735 1.0000
8.500 0.9148 0.03128 0.02065 -0.0145 0.2441 1.0000
8.750 0.9214 0.03270 0.02198 -0.0121 0.2181 1.0000
9.000 0.9296 0.03408 0.02330 -0.0100 0.1984 1.0000
9.250 0.9392 0.03546 0.02462 -0.0081 0.1844 1.0000
9.500 0.9505 0.03684 0.02600 -0.0065 0.1728 1.0000
9.750 0.9628 0.03824 0.02735 -0.0050 0.1628 1.0000
10.000 0.9772 0.03962 0.02872 -0.0037 0.1535 1.0000
10.250 0.9973 0.04096 0.03019 -0.0028 0.1452 1.0000
10.500 1.0188 0.04235 0.03150 -0.0021 0.1367 1.0000
10.750 1.0350 0.04395 0.03340 -0.0011 0.1283 1.0000
11.000 1.0490 0.04561 0.03507 0.0000 0.1195 1.0000
11.250 1.0515 0.04739 0.03697 0.0017 0.1107 1.0000
11.500 1.0536 0.04936 0.03909 0.0032 0.1026 1.0000
11.750 1.0569 0.05127 0.04103 0.0045 0.0954 1.0000
12.000 1.0595 0.05364 0.04365 0.0058 0.0887 1.0000
12.250 1.0619 0.05589 0.04600 0.0069 0.0827 1.0000
12.500 1.0644 0.05854 0.04882 0.0078 0.0775 1.0000
12.750 1.0614 0.06155 0.05212 0.0086 0.0727 1.0000
13.000 1.0658 0.06392 0.05446 0.0093 0.0686 1.0000
13.250 1.0591 0.06781 0.05867 0.0096 0.0657 1.0000
13.500 1.0476 0.07220 0.06340 0.0095 0.0632 1.0000
13.750 1.0361 0.07664 0.06807 0.0089 0.0612 1.0000
14.000 1.0261 0.08096 0.07254 0.0080 0.0593 1.0000
14.250 1.0209 0.08471 0.07636 0.0072 0.0575 1.0000
14.500 1.0116 0.08938 0.08111 0.0059 0.0562 1.0000
14.750 0.9863 0.09683 0.08883 0.0025 0.0560 1.0000
15.000 0.9607 0.10510 0.09731 -0.0017 0.0560 1.0000
15.250 0.9326 0.11471 0.10709 -0.0072 0.0563 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 31 AIRFOIL (raf31-il)