GOE 510 AIRFOIL (goe510-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 510 AIRFOIL (goe510-il) Reynolds number: 200,000 Max Cl/Cd: 61.59 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe510-il-200000-n5.txt Download as CSV file: xf-goe510-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 510 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.1744 0.11622 0.11238 -0.0673 0.9760 0.0493
-10.500 -0.1723 0.11196 0.10811 -0.0727 0.9730 0.0496
-10.250 -0.1647 0.10743 0.10357 -0.0777 0.9709 0.0498
-10.000 -0.1611 0.10395 0.10009 -0.0793 0.9653 0.0498
-9.750 -0.1515 0.10002 0.09616 -0.0823 0.9617 0.0498
-9.500 -0.1387 0.09602 0.09214 -0.0858 0.9591 0.0498
-9.000 -0.1143 0.08981 0.08594 -0.0879 0.9496 0.0494
-8.750 -0.0911 0.08753 0.08366 -0.0881 0.9468 0.0476
-8.500 -0.0783 0.08337 0.07949 -0.0926 0.9437 0.0475
-8.250 -0.0792 0.08012 0.07625 -0.0937 0.9343 0.0476
-8.000 -0.0693 0.07593 0.07205 -0.0978 0.9297 0.0466
-7.750 -0.0765 0.07197 0.06810 -0.0997 0.9187 0.0457
-7.500 -0.0844 0.06700 0.06308 -0.1028 0.9095 0.0449
-7.250 -0.0987 0.05927 0.05522 -0.1071 0.8983 0.0438
-7.000 -0.1024 0.05511 0.05096 -0.1072 0.8870 0.0439
-6.750 -0.0820 0.05197 0.04773 -0.1098 0.8821 0.0444
-6.500 -0.0768 0.04921 0.04488 -0.1089 0.8712 0.0450
-6.250 -0.0610 0.04341 0.03883 -0.1122 0.8656 0.0457
-6.000 -0.0830 0.03358 0.02828 -0.1094 0.8527 0.0458
-5.750 -0.0732 0.02997 0.02419 -0.1078 0.8438 0.0462
-5.500 -0.0511 0.02730 0.02107 -0.1079 0.8359 0.0468
-5.250 -0.0309 0.02544 0.01883 -0.1069 0.8269 0.0475
-5.000 -0.0034 0.02384 0.01683 -0.1072 0.8194 0.0487
-4.750 0.0204 0.02270 0.01531 -0.1065 0.8110 0.0495
-4.500 0.0484 0.02156 0.01387 -0.1067 0.8028 0.0501
-4.250 0.0740 0.02054 0.01272 -0.1064 0.7934 0.0506
-4.000 0.1048 0.01971 0.01175 -0.1071 0.7841 0.0513
-3.750 0.1286 0.01909 0.01102 -0.1062 0.7731 0.0519
-3.500 0.1578 0.01848 0.01026 -0.1064 0.7632 0.0527
-3.250 0.1833 0.01799 0.00965 -0.1059 0.7519 0.0537
-3.000 0.2084 0.01757 0.00911 -0.1052 0.7400 0.0550
-2.750 0.2351 0.01715 0.00855 -0.1048 0.7277 0.0562
-2.500 0.2619 0.01677 0.00801 -0.1044 0.7150 0.0571
-2.250 0.2859 0.01645 0.00757 -0.1034 0.7016 0.0577
-2.000 0.3103 0.01601 0.00707 -0.1026 0.6889 0.0586
-1.750 0.3344 0.01564 0.00666 -0.1018 0.6761 0.0596
-1.500 0.3582 0.01539 0.00635 -0.1009 0.6631 0.0607
-1.250 0.3806 0.01522 0.00611 -0.0996 0.6495 0.0625
-1.000 0.4024 0.01508 0.00591 -0.0983 0.6366 0.0642
-0.750 0.4243 0.01496 0.00572 -0.0970 0.6243 0.0657
-0.500 0.4458 0.01488 0.00554 -0.0955 0.6118 0.0668
-0.250 0.4658 0.01479 0.00539 -0.0938 0.5986 0.0681
0.000 0.4853 0.01468 0.00525 -0.0920 0.5856 0.0700
0.250 0.5048 0.01464 0.00515 -0.0902 0.5725 0.0725
0.500 0.5239 0.01463 0.00507 -0.0883 0.5595 0.0756
1.000 0.5586 0.01459 0.00495 -0.0838 0.5318 0.0833
1.500 0.5907 0.01448 0.00493 -0.0789 0.5068 0.1263
1.750 0.6061 0.01444 0.00506 -0.0764 0.4953 0.1987
2.000 0.6235 0.01454 0.00520 -0.0743 0.4844 0.2361
2.250 0.6408 0.01469 0.00532 -0.0722 0.4746 0.2580
2.500 0.6595 0.01484 0.00544 -0.0704 0.4658 0.2754
2.750 0.6777 0.01501 0.00555 -0.0685 0.4578 0.2887
3.000 0.6968 0.01516 0.00568 -0.0667 0.4501 0.3021
3.500 0.7344 0.01544 0.00596 -0.0633 0.4367 0.3349
3.750 0.7533 0.01558 0.00609 -0.0616 0.4306 0.3519
4.000 0.7720 0.01571 0.00624 -0.0599 0.4252 0.3794
4.500 0.9542 0.01596 0.00768 -0.0873 0.4075 0.9929
4.750 0.9955 0.01628 0.00799 -0.0905 0.4012 1.0000
5.000 1.0094 0.01648 0.00815 -0.0878 0.3966 1.0000
5.250 1.0236 0.01672 0.00832 -0.0853 0.3925 1.0000
5.500 1.0387 0.01692 0.00852 -0.0829 0.3883 1.0000
5.750 1.0534 0.01712 0.00872 -0.0804 0.3836 1.0000
6.000 1.0677 0.01735 0.00893 -0.0779 0.3787 1.0000
6.250 1.0829 0.01762 0.00916 -0.0757 0.3748 1.0000
6.500 1.0983 0.01786 0.00940 -0.0734 0.3703 1.0000
6.750 1.1141 0.01809 0.00967 -0.0713 0.3659 1.0000
7.000 1.1293 0.01837 0.00994 -0.0691 0.3616 1.0000
7.250 1.1443 0.01869 0.01022 -0.0669 0.3574 1.0000
7.500 1.1599 0.01896 0.01054 -0.0649 0.3526 1.0000
7.750 1.1754 0.01926 0.01087 -0.0628 0.3479 1.0000
8.000 1.1903 0.01960 0.01121 -0.0607 0.3435 1.0000
8.250 1.2056 0.01996 0.01156 -0.0587 0.3396 1.0000
8.500 1.2212 0.02029 0.01197 -0.0568 0.3345 1.0000
8.750 1.2347 0.02067 0.01238 -0.0546 0.3290 1.0000
9.000 1.2461 0.02112 0.01279 -0.0521 0.3227 1.0000
9.250 1.2590 0.02154 0.01328 -0.0499 0.3156 1.0000
9.500 1.2683 0.02208 0.01379 -0.0473 0.3081 1.0000
9.750 1.2797 0.02259 0.01436 -0.0450 0.2994 1.0000
10.000 1.2879 0.02325 0.01499 -0.0424 0.2908 1.0000
10.250 1.2976 0.02390 0.01567 -0.0401 0.2807 1.0000
10.500 1.3061 0.02466 0.01643 -0.0378 0.2714 1.0000
10.750 1.3145 0.02548 0.01725 -0.0355 0.2628 1.0000
11.000 1.3242 0.02629 0.01809 -0.0335 0.2545 1.0000
11.250 1.3294 0.02736 0.01913 -0.0312 0.2443 1.0000
11.500 1.3365 0.02841 0.02019 -0.0291 0.2336 1.0000
11.750 1.3426 0.02956 0.02136 -0.0271 0.2239 1.0000
12.000 1.3460 0.03094 0.02272 -0.0250 0.2138 1.0000
12.250 1.3507 0.03232 0.02411 -0.0231 0.2027 1.0000
12.500 1.3537 0.03388 0.02568 -0.0213 0.1911 1.0000
12.750 1.3518 0.03588 0.02763 -0.0193 0.1755 1.0000
13.000 1.3477 0.03819 0.02989 -0.0174 0.1590 1.0000
13.250 1.3387 0.04105 0.03266 -0.0155 0.1393 1.0000
13.500 1.3287 0.04417 0.03572 -0.0139 0.1220 1.0000
13.750 1.3188 0.04748 0.03899 -0.0127 0.1076 1.0000
14.000 1.3068 0.05114 0.04262 -0.0116 0.0938 1.0000
14.250 1.2923 0.05525 0.04669 -0.0108 0.0773 1.0000
14.500 1.2659 0.06091 0.05225 -0.0104 0.0434 1.0000
14.750 1.2415 0.06665 0.05795 -0.0103 0.0305 1.0000
15.000 1.2314 0.07086 0.06226 -0.0105 0.0281 1.0000
15.250 1.2223 0.07508 0.06658 -0.0108 0.0262 1.0000
15.500 1.2159 0.07900 0.07062 -0.0112 0.0251 1.0000
15.750 1.2090 0.08308 0.07482 -0.0117 0.0239 1.0000
16.000 1.2029 0.08709 0.07895 -0.0123 0.0233 1.0000
16.250 1.1956 0.09135 0.08332 -0.0130 0.0226 1.0000
16.500 1.1872 0.09577 0.08786 -0.0139 0.0219 1.0000
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Polar data table (+)
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