GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 611 AIRFOIL (goe611-il) Reynolds number: 500,000 Max Cl/Cd: 106.56 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe611-il-500000.txt Download as CSV file: xf-goe611-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 611 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.0620 0.08936 0.08722 -0.0784 0.9416 0.0189
-9.750 -0.0622 0.08553 0.08336 -0.0803 0.9277 0.0189
-9.500 -0.0626 0.08175 0.07954 -0.0819 0.9137 0.0190
-9.250 -0.0599 0.07776 0.07551 -0.0817 0.9010 0.0191
-9.000 -0.0514 0.07495 0.07265 -0.0806 0.8893 0.0194
-8.750 -0.0461 0.07204 0.06969 -0.0806 0.8790 0.0196
-8.500 -0.0425 0.06900 0.06659 -0.0810 0.8698 0.0198
-8.250 -0.0394 0.06583 0.06340 -0.0819 0.8607 0.0201
-8.000 -0.0370 0.06259 0.06012 -0.0829 0.8529 0.0204
-7.750 -0.0352 0.05922 0.05673 -0.0843 0.8450 0.0208
-7.500 -0.1002 0.06903 0.06650 -0.0891 0.8603 0.0199
-7.250 -0.0889 0.06556 0.06297 -0.0927 0.8516 0.0203
-7.000 -0.0759 0.06187 0.05920 -0.0966 0.8433 0.0208
-6.750 -0.0608 0.05795 0.05519 -0.1006 0.8359 0.0215
-6.500 -0.0348 0.05146 0.04840 -0.1084 0.8284 0.0233
-6.250 -0.0152 0.04688 0.04353 -0.1108 0.8221 0.0234
-6.000 -0.0028 0.04198 0.03850 -0.1120 0.8154 0.0237
-5.750 0.0145 0.03986 0.03631 -0.1123 0.8095 0.0241
-5.500 0.0342 0.03808 0.03448 -0.1128 0.8028 0.0247
-5.250 0.0559 0.03595 0.03220 -0.1132 0.7965 0.0258
-5.000 0.0866 0.03323 0.02891 -0.1131 0.7907 0.0288
-4.250 0.1521 0.02534 0.02058 -0.1137 0.7714 0.0312
-4.000 0.1786 0.02406 0.01906 -0.1135 0.7650 0.0341
-3.750 0.2047 0.02166 0.01622 -0.1130 0.7584 0.0372
-3.500 0.2303 0.02044 0.01494 -0.1132 0.7512 0.0386
-3.250 0.2573 0.01952 0.01387 -0.1131 0.7440 0.0415
-3.000 0.2847 0.01822 0.01222 -0.1128 0.7362 0.0468
-2.750 0.3112 0.01734 0.01133 -0.1130 0.7283 0.0499
-2.500 0.3389 0.01667 0.01037 -0.1129 0.7199 0.0583
-2.250 0.3658 0.01574 0.00946 -0.1130 0.7106 0.0621
-2.000 0.3930 0.01514 0.00867 -0.1129 0.7006 0.0728
-1.750 0.4202 0.01468 0.00817 -0.1129 0.6892 0.0802
-1.500 0.4519 0.01314 0.00614 -0.1115 0.6783 0.0453
-1.250 0.4790 0.01231 0.00526 -0.1112 0.6669 0.0437
-1.000 0.5062 0.01189 0.00476 -0.1110 0.6556 0.0439
-0.750 0.5333 0.01159 0.00437 -0.1108 0.6438 0.0443
-0.500 0.5601 0.01123 0.00396 -0.1106 0.6312 0.0443
-0.250 0.5870 0.01100 0.00367 -0.1104 0.6186 0.0446
0.000 0.6139 0.01082 0.00343 -0.1103 0.6069 0.0450
0.250 0.6407 0.01061 0.00314 -0.1101 0.5961 0.0458
0.500 0.6678 0.01048 0.00296 -0.1101 0.5851 0.0483
0.750 0.6951 0.01044 0.00287 -0.1100 0.5747 0.0501
1.000 0.7222 0.01046 0.00282 -0.1100 0.5649 0.0525
1.250 0.7493 0.01050 0.00279 -0.1099 0.5551 0.0557
1.500 0.7765 0.01049 0.00281 -0.1099 0.5458 0.0818
1.750 0.8023 0.00858 0.00306 -0.1099 0.5373 1.0000
2.000 0.8288 0.00870 0.00310 -0.1097 0.5283 1.0000
2.250 0.8549 0.00886 0.00317 -0.1095 0.5200 1.0000
2.500 0.8812 0.00901 0.00325 -0.1093 0.5115 1.0000
2.750 0.9075 0.00916 0.00334 -0.1092 0.5036 1.0000
3.000 0.9334 0.00933 0.00344 -0.1089 0.4947 1.0000
3.250 0.9597 0.00948 0.00355 -0.1088 0.4860 1.0000
3.500 0.9852 0.00967 0.00367 -0.1085 0.4777 1.0000
3.750 1.0116 0.00981 0.00379 -0.1084 0.4691 1.0000
4.000 1.0370 0.01000 0.00393 -0.1081 0.4610 1.0000
4.250 1.0629 0.01016 0.00408 -0.1080 0.4522 1.0000
4.500 1.0883 0.01035 0.00423 -0.1077 0.4440 1.0000
4.750 1.1136 0.01054 0.00440 -0.1074 0.4352 1.0000
5.000 1.1388 0.01072 0.00458 -0.1072 0.4266 1.0000
5.500 1.1881 0.01115 0.00495 -0.1064 0.4079 1.0000
5.750 1.2121 0.01139 0.00517 -0.1059 0.3978 1.0000
6.000 1.2354 0.01165 0.00541 -0.1054 0.3880 1.0000
6.250 1.2588 0.01191 0.00564 -0.1048 0.3759 1.0000
6.500 1.2804 0.01224 0.00592 -0.1039 0.3603 1.0000
6.750 1.3006 0.01263 0.00624 -0.1028 0.3428 1.0000
7.000 1.3205 0.01303 0.00657 -0.1017 0.3239 1.0000
7.250 1.3383 0.01352 0.00697 -0.1002 0.3028 1.0000
7.500 1.3548 0.01405 0.00740 -0.0985 0.2822 1.0000
7.750 1.3664 0.01473 0.00795 -0.0959 0.2556 1.0000
8.000 1.3750 0.01557 0.00861 -0.0929 0.2242 1.0000
8.250 1.3770 0.01681 0.00959 -0.0890 0.1786 1.0000
8.500 1.3798 0.01810 0.01066 -0.0854 0.1455 1.0000
8.750 1.3900 0.01903 0.01152 -0.0831 0.1345 1.0000
9.000 1.4024 0.01984 0.01234 -0.0813 0.1277 1.0000
9.250 1.4123 0.02085 0.01333 -0.0792 0.1204 1.0000
9.500 1.4265 0.02161 0.01413 -0.0777 0.1148 1.0000
9.750 1.4356 0.02273 0.01523 -0.0758 0.1078 1.0000
10.000 1.4505 0.02351 0.01603 -0.0746 0.0997 1.0000
10.250 1.4613 0.02461 0.01709 -0.0730 0.0862 1.0000
10.500 1.4404 0.02811 0.02023 -0.0686 0.0315 1.0000
10.750 1.4383 0.03040 0.02250 -0.0662 0.0193 1.0000
11.000 1.4452 0.03203 0.02422 -0.0648 0.0176 1.0000
11.250 1.4507 0.03383 0.02612 -0.0634 0.0166 1.0000
11.500 1.4546 0.03580 0.02820 -0.0620 0.0157 1.0000
11.750 1.4564 0.03802 0.03055 -0.0606 0.0150 1.0000
12.000 1.4596 0.04018 0.03283 -0.0594 0.0146 1.0000
12.250 1.4621 0.04246 0.03522 -0.0584 0.0143 1.0000
12.500 1.4630 0.04498 0.03785 -0.0574 0.0138 1.0000
12.750 1.4625 0.04774 0.04072 -0.0565 0.0134 1.0000
13.000 1.4603 0.05077 0.04386 -0.0557 0.0130 1.0000
13.250 1.4564 0.05411 0.04731 -0.0551 0.0126 1.0000
13.500 1.4508 0.05775 0.05106 -0.0547 0.0124 1.0000
13.750 1.4433 0.06175 0.05518 -0.0544 0.0121 1.0000
14.000 1.4346 0.06602 0.05958 -0.0543 0.0119 1.0000
14.250 1.4232 0.07073 0.06440 -0.0544 0.0118 1.0000
14.500 1.4105 0.07572 0.06951 -0.0546 0.0116 1.0000
14.750 1.3969 0.08092 0.07483 -0.0550 0.0115 1.0000
15.000 1.3851 0.08590 0.07991 -0.0555 0.0114 1.0000
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Polar data table (+)
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