GOE 599 AIRFOIL (goe599-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 599 AIRFOIL (goe599-il) Reynolds number: 200,000 Max Cl/Cd: 51.23 at α=2.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe599-il-200000-n5.txt Download as CSV file: xf-goe599-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 599 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 -0.6283 0.08198 0.07841 -0.0324 1.0000 0.0135
-10.500 -0.6944 0.05868 0.05498 -0.0509 1.0000 0.0125
-10.250 -0.7274 0.05167 0.04776 -0.0553 1.0000 0.0123
-10.000 -0.7522 0.04738 0.04322 -0.0536 1.0000 0.0122
-9.750 -0.7663 0.04348 0.03902 -0.0515 1.0000 0.0123
-9.500 -0.7730 0.03994 0.03515 -0.0492 1.0000 0.0125
-9.250 -0.7735 0.03678 0.03163 -0.0470 1.0000 0.0127
-9.000 -0.7691 0.03393 0.02842 -0.0448 1.0000 0.0130
-8.750 -0.7609 0.03137 0.02550 -0.0427 1.0000 0.0135
-8.500 -0.7490 0.02924 0.02298 -0.0409 1.0000 0.0141
-8.250 -0.7339 0.02759 0.02101 -0.0392 1.0000 0.0150
-8.000 -0.7189 0.02567 0.01880 -0.0376 1.0000 0.0158
-7.750 -0.7024 0.02415 0.01714 -0.0362 1.0000 0.0166
-7.500 -0.6838 0.02304 0.01589 -0.0350 1.0000 0.0175
-7.250 -0.6646 0.02198 0.01470 -0.0337 1.0000 0.0185
-7.000 -0.6447 0.02101 0.01358 -0.0326 1.0000 0.0200
-6.750 -0.6239 0.02024 0.01265 -0.0314 1.0000 0.0218
-6.500 -0.6039 0.01918 0.01145 -0.0303 1.0000 0.0235
-6.250 -0.5834 0.01832 0.01054 -0.0293 1.0000 0.0255
-6.000 -0.5539 0.01764 0.00978 -0.0301 0.9982 0.0288
-5.750 -0.5208 0.01695 0.00896 -0.0316 0.9956 0.0324
-5.500 -0.4888 0.01611 0.00807 -0.0330 0.9927 0.0371
-5.250 -0.4555 0.01562 0.00744 -0.0344 0.9894 0.0421
-5.000 -0.4220 0.01492 0.00673 -0.0361 0.9866 0.0491
-4.750 -0.3893 0.01442 0.00619 -0.0374 0.9830 0.0579
-4.500 -0.3566 0.01391 0.00581 -0.0388 0.9791 0.0795
-4.250 -0.3216 0.01354 0.00543 -0.0406 0.9761 0.1087
-4.000 -0.2902 0.01313 0.00506 -0.0417 0.9714 0.1296
-3.750 -0.2575 0.01262 0.00476 -0.0432 0.9672 0.1723
-3.500 -0.2228 0.01214 0.00452 -0.0451 0.9642 0.2431
-3.250 -0.1927 0.01179 0.00433 -0.0458 0.9583 0.2911
-3.000 -0.1595 0.01145 0.00414 -0.0472 0.9539 0.3393
-2.750 -0.1252 0.01104 0.00397 -0.0489 0.9507 0.4047
-2.500 -0.0976 0.01071 0.00387 -0.0490 0.9431 0.4664
-2.250 -0.0654 0.01045 0.00376 -0.0499 0.9380 0.5175
-2.000 -0.0356 0.01024 0.00365 -0.0503 0.9313 0.5601
-1.750 -0.0052 0.01006 0.00355 -0.0508 0.9244 0.5936
-1.500 0.0249 0.00988 0.00345 -0.0511 0.9171 0.6237
-1.250 0.0551 0.00967 0.00333 -0.0515 0.9084 0.6553
-1.000 0.0844 0.00943 0.00320 -0.0514 0.8957 0.6913
-0.750 0.1139 0.00919 0.00307 -0.0513 0.8821 0.7315
-0.500 0.1429 0.00899 0.00298 -0.0511 0.8696 0.7725
-0.250 0.1713 0.00885 0.00295 -0.0507 0.8593 0.8158
0.000 0.2009 0.00876 0.00293 -0.0505 0.8497 0.8604
0.250 0.2332 0.00872 0.00295 -0.0510 0.8392 0.9043
0.500 0.2711 0.00872 0.00295 -0.0528 0.8285 0.9409
0.750 0.3107 0.00873 0.00289 -0.0549 0.8088 0.9687
1.000 0.3507 0.00878 0.00282 -0.0572 0.7790 0.9893
1.250 0.3867 0.00886 0.00278 -0.0588 0.7477 1.0000
1.500 0.4103 0.00894 0.00278 -0.0579 0.7205 1.0000
1.750 0.4340 0.00904 0.00278 -0.0570 0.6926 1.0000
2.000 0.4576 0.00918 0.00281 -0.0560 0.6613 1.0000
2.250 0.4804 0.00938 0.00284 -0.0549 0.6198 1.0000
2.500 0.5005 0.00977 0.00289 -0.0532 0.5506 1.0000
2.750 0.5200 0.01028 0.00304 -0.0516 0.4755 1.0000
3.000 0.5399 0.01083 0.00325 -0.0502 0.4080 1.0000
3.250 0.5607 0.01138 0.00351 -0.0490 0.3450 1.0000
3.500 0.5805 0.01209 0.00383 -0.0477 0.2567 1.0000
3.750 0.5965 0.01335 0.00439 -0.0461 0.1197 1.0000
4.000 0.6188 0.01393 0.00477 -0.0452 0.0830 1.0000
4.250 0.6428 0.01433 0.00514 -0.0446 0.0755 1.0000
4.500 0.6668 0.01472 0.00557 -0.0440 0.0715 1.0000
4.750 0.6903 0.01517 0.00604 -0.0433 0.0683 1.0000
5.000 0.7137 0.01563 0.00654 -0.0425 0.0659 1.0000
5.250 0.7372 0.01606 0.00705 -0.0418 0.0638 1.0000
5.500 0.7602 0.01655 0.00763 -0.0410 0.0622 1.0000
5.750 0.7827 0.01709 0.00824 -0.0402 0.0607 1.0000
6.000 0.8049 0.01769 0.00890 -0.0393 0.0592 1.0000
6.250 0.8267 0.01835 0.00961 -0.0383 0.0578 1.0000
6.500 0.8477 0.01912 0.01041 -0.0373 0.0557 1.0000
6.750 0.8679 0.02008 0.01142 -0.0362 0.0522 1.0000
7.000 0.8919 0.02038 0.01186 -0.0357 0.0487 1.0000
7.250 0.9145 0.02089 0.01245 -0.0350 0.0452 1.0000
7.500 0.9352 0.02162 0.01318 -0.0341 0.0418 1.0000
7.750 0.9578 0.02212 0.01384 -0.0334 0.0384 1.0000
8.000 0.9809 0.02242 0.01429 -0.0327 0.0341 1.0000
8.250 1.0034 0.02279 0.01475 -0.0320 0.0299 1.0000
8.500 1.0259 0.02318 0.01522 -0.0313 0.0256 1.0000
8.750 1.0458 0.02394 0.01606 -0.0302 0.0225 1.0000
9.000 1.0660 0.02460 0.01679 -0.0292 0.0197 1.0000
9.250 1.0840 0.02549 0.01768 -0.0279 0.0175 1.0000
9.500 1.1003 0.02673 0.01911 -0.0264 0.0162 1.0000
9.750 1.1159 0.02804 0.02059 -0.0247 0.0152 1.0000
10.000 1.1306 0.02938 0.02210 -0.0230 0.0143 1.0000
10.250 1.1440 0.03074 0.02365 -0.0213 0.0135 1.0000
10.500 1.1540 0.03224 0.02529 -0.0191 0.0128 1.0000
10.750 1.1580 0.03440 0.02764 -0.0163 0.0121 1.0000
11.000 1.1657 0.03616 0.02968 -0.0140 0.0116 1.0000
11.250 1.1701 0.03823 0.03203 -0.0116 0.0112 1.0000
11.500 1.1707 0.04062 0.03470 -0.0090 0.0108 1.0000
11.750 1.1679 0.04330 0.03766 -0.0065 0.0106 1.0000
12.000 1.1619 0.04631 0.04096 -0.0043 0.0104 1.0000
12.250 1.1526 0.04973 0.04466 -0.0026 0.0102 1.0000
12.500 1.1400 0.05368 0.04889 -0.0016 0.0101 1.0000
12.750 1.1241 0.05833 0.05382 -0.0015 0.0101 1.0000
13.000 1.1047 0.06389 0.05964 -0.0026 0.0100 1.0000
13.250 1.0817 0.07063 0.06664 -0.0053 0.0101 1.0000
13.500 1.0544 0.07913 0.07540 -0.0100 0.0102 1.0000
13.750 1.0208 0.09040 0.08690 -0.0174 0.0104 1.0000
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Polar data table (+)
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