GOE 566 AIRFOIL (goe566-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 566 AIRFOIL (goe566-il) Reynolds number: 100,000 Max Cl/Cd: 50.27 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe566-il-100000-n5.txt Download as CSV file: xf-goe566-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 566 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4301 0.11242 0.10711 -0.0218 1.0000 0.0530
-10.000 -0.4379 0.10951 0.10428 -0.0261 1.0000 0.0566
-9.750 -0.4432 0.10614 0.10098 -0.0297 1.0000 0.0571
-9.250 -0.4312 0.09569 0.09052 -0.0294 1.0000 0.0427
-9.000 -0.4298 0.09196 0.08684 -0.0302 1.0000 0.0413
-8.750 -0.4313 0.08807 0.08301 -0.0315 1.0000 0.0404
-8.500 -0.4351 0.08417 0.07917 -0.0331 1.0000 0.0395
-8.250 -0.4421 0.08039 0.07548 -0.0346 1.0000 0.0394
-8.000 -0.4516 0.07661 0.07178 -0.0363 1.0000 0.0397
-7.750 -0.4578 0.07246 0.06766 -0.0383 1.0000 0.0399
-7.500 -0.4634 0.06830 0.06351 -0.0396 1.0000 0.0399
-7.250 -0.4687 0.06416 0.05934 -0.0402 1.0000 0.0403
-7.000 -0.4732 0.06013 0.05523 -0.0399 1.0000 0.0405
-6.750 -0.4765 0.05611 0.05109 -0.0391 1.0000 0.0403
-6.500 -0.4778 0.05202 0.04682 -0.0380 1.0000 0.0403
-6.250 -0.4764 0.04790 0.04245 -0.0368 1.0000 0.0404
-6.000 -0.4720 0.04386 0.03807 -0.0354 1.0000 0.0418
-5.750 -0.4456 0.03836 0.03182 -0.0381 0.9942 0.0439
-5.500 -0.4182 0.03445 0.02721 -0.0396 0.9884 0.0446
-5.250 -0.3910 0.03151 0.02401 -0.0412 0.9834 0.0468
-5.000 -0.3616 0.02991 0.02220 -0.0425 0.9780 0.0492
-4.750 -0.3306 0.02767 0.01951 -0.0437 0.9734 0.0503
-4.500 -0.3011 0.02580 0.01724 -0.0443 0.9677 0.0515
-4.250 -0.2674 0.02438 0.01545 -0.0457 0.9636 0.0542
-4.000 -0.2374 0.02318 0.01391 -0.0461 0.9577 0.0562
-3.750 -0.2047 0.02202 0.01248 -0.0471 0.9529 0.0570
-3.500 -0.1695 0.02081 0.01114 -0.0486 0.9494 0.0582
-3.250 -0.1411 0.01988 0.01017 -0.0488 0.9430 0.0598
-3.000 -0.1083 0.01914 0.00941 -0.0499 0.9380 0.0620
-2.750 -0.0723 0.01856 0.00880 -0.0516 0.9345 0.0665
-2.500 -0.0448 0.01808 0.00824 -0.0515 0.9271 0.0700
-2.250 -0.0125 0.01748 0.00766 -0.0524 0.9220 0.0734
-2.000 0.0182 0.01706 0.00724 -0.0530 0.9162 0.0788
-1.750 0.0474 0.01668 0.00686 -0.0532 0.9093 0.0876
-1.500 0.0814 0.01619 0.00649 -0.0543 0.9049 0.1163
-1.250 0.1061 0.01566 0.00634 -0.0539 0.8964 0.2041
-1.000 0.1294 0.01381 0.00635 -0.0531 0.8913 0.6795
-0.750 0.2390 0.01318 0.00620 -0.0683 0.8967 1.0000
-0.500 0.2704 0.01305 0.00591 -0.0686 0.8843 1.0000
-0.250 0.2994 0.01291 0.00563 -0.0684 0.8698 1.0000
0.000 0.3259 0.01279 0.00537 -0.0676 0.8536 1.0000
0.250 0.3512 0.01269 0.00516 -0.0666 0.8369 1.0000
0.500 0.3760 0.01263 0.00498 -0.0655 0.8206 1.0000
0.750 0.4003 0.01261 0.00488 -0.0644 0.8057 1.0000
1.000 0.4245 0.01263 0.00483 -0.0635 0.7922 1.0000
1.250 0.4489 0.01266 0.00480 -0.0625 0.7785 1.0000
1.500 0.4733 0.01269 0.00478 -0.0615 0.7638 1.0000
1.750 0.4976 0.01273 0.00477 -0.0605 0.7482 1.0000
2.000 0.5219 0.01278 0.00477 -0.0594 0.7317 1.0000
2.250 0.5464 0.01284 0.00480 -0.0584 0.7146 1.0000
2.500 0.5708 0.01291 0.00482 -0.0574 0.6960 1.0000
2.750 0.5945 0.01301 0.00488 -0.0562 0.6742 1.0000
3.000 0.6185 0.01312 0.00493 -0.0551 0.6504 1.0000
3.250 0.6419 0.01328 0.00500 -0.0539 0.6228 1.0000
3.500 0.6646 0.01347 0.00511 -0.0526 0.5893 1.0000
3.750 0.6865 0.01372 0.00524 -0.0511 0.5475 1.0000
4.000 0.7073 0.01407 0.00537 -0.0495 0.4960 1.0000
4.250 0.7270 0.01455 0.00556 -0.0478 0.4462 1.0000
4.500 0.7470 0.01508 0.00586 -0.0463 0.4073 1.0000
4.750 0.7679 0.01558 0.00625 -0.0451 0.3780 1.0000
5.000 0.7893 0.01607 0.00665 -0.0439 0.3550 1.0000
5.250 0.8110 0.01655 0.00707 -0.0428 0.3348 1.0000
5.500 0.8329 0.01702 0.00753 -0.0418 0.3171 1.0000
5.750 0.8548 0.01750 0.00800 -0.0408 0.3019 1.0000
6.000 0.8769 0.01798 0.00849 -0.0398 0.2884 1.0000
6.250 0.8989 0.01847 0.00900 -0.0389 0.2762 1.0000
6.500 0.9203 0.01899 0.00955 -0.0379 0.2623 1.0000
6.750 0.9413 0.01950 0.01010 -0.0368 0.2464 1.0000
7.000 0.9620 0.02002 0.01067 -0.0357 0.2290 1.0000
7.250 0.9821 0.02055 0.01124 -0.0345 0.2096 1.0000
7.500 1.0016 0.02110 0.01182 -0.0333 0.1897 1.0000
7.750 1.0223 0.02161 0.01248 -0.0323 0.1661 1.0000
8.000 1.0405 0.02230 0.01312 -0.0310 0.1368 1.0000
8.250 1.0557 0.02333 0.01394 -0.0294 0.1092 1.0000
8.500 1.0696 0.02455 0.01508 -0.0276 0.0899 1.0000
8.750 1.0827 0.02585 0.01639 -0.0257 0.0742 1.0000
9.000 1.0960 0.02707 0.01772 -0.0238 0.0618 1.0000
9.250 1.1091 0.02824 0.01893 -0.0220 0.0528 1.0000
9.500 1.1194 0.02955 0.02022 -0.0199 0.0471 1.0000
9.750 1.1289 0.03082 0.02163 -0.0176 0.0426 1.0000
10.000 1.1342 0.03224 0.02309 -0.0149 0.0393 1.0000
10.250 1.1371 0.03389 0.02483 -0.0121 0.0370 1.0000
10.500 1.1423 0.03552 0.02663 -0.0097 0.0350 1.0000
10.750 1.1466 0.03723 0.02848 -0.0075 0.0328 1.0000
11.000 1.1496 0.03910 0.03043 -0.0055 0.0311 1.0000
11.250 1.1506 0.04137 0.03278 -0.0036 0.0296 1.0000
11.500 1.1554 0.04354 0.03513 -0.0020 0.0285 1.0000
11.750 1.1600 0.04585 0.03768 -0.0005 0.0275 1.0000
12.000 1.1624 0.04832 0.04040 0.0008 0.0263 1.0000
12.250 1.1636 0.05098 0.04327 0.0019 0.0256 1.0000
12.500 1.1617 0.05384 0.04633 0.0026 0.0246 1.0000
12.750 1.1583 0.05682 0.04946 0.0030 0.0237 1.0000
13.000 1.1542 0.06009 0.05288 0.0031 0.0231 1.0000
13.250 1.1488 0.06370 0.05663 0.0029 0.0225 1.0000
13.500 1.1414 0.06780 0.06086 0.0023 0.0220 1.0000
13.750 1.1317 0.07247 0.06572 0.0013 0.0218 1.0000
14.000 1.1192 0.07764 0.07108 -0.0006 0.0217 1.0000
14.250 1.1044 0.08341 0.07709 -0.0032 0.0216 1.0000
14.500 1.0872 0.09000 0.08390 -0.0067 0.0216 1.0000
14.750 1.0705 0.09696 0.09106 -0.0107 0.0216 1.0000
15.000 1.0516 0.10495 0.09924 -0.0156 0.0217 1.0000
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