GOE 5K AIRFOIL (goe05k-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 5K AIRFOIL (goe05k-il) Reynolds number: 50,000 Max Cl/Cd: 20.52 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe05k-il-50000.txt Download as CSV file: xf-goe05k-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 5K 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5621 0.10889 0.10220 -0.0093 1.0000 0.1390
-8.000 -0.5477 0.10313 0.09640 -0.0073 1.0000 0.1507
-7.750 -0.5586 0.10200 0.09541 -0.0103 1.0000 0.1537
-7.500 -0.5483 0.09712 0.09055 -0.0081 1.0000 0.1658
-7.250 -0.5454 0.09347 0.08697 -0.0084 1.0000 0.1748
-7.000 -0.5497 0.09175 0.08534 -0.0129 1.0000 0.1833
-6.750 -0.5449 0.08800 0.08165 -0.0125 1.0000 0.1980
-6.500 -0.5415 0.08479 0.07848 -0.0134 1.0000 0.2133
-6.250 -0.5335 0.08027 0.07403 -0.0100 1.0000 0.2324
-6.000 -0.5310 0.07754 0.07134 -0.0107 1.0000 0.2573
-5.750 -0.5252 0.07392 0.06778 -0.0076 1.0000 0.2859
-5.500 -0.5222 0.07068 0.06460 -0.0054 1.0000 0.3197
-5.250 -0.5184 0.06746 0.06147 -0.0001 1.0000 0.3650
-5.000 -0.2326 0.04542 0.03841 0.0007 1.0000 0.9683
-4.750 -0.2491 0.04484 0.03803 0.0060 1.0000 0.9441
-4.500 -0.2852 0.04541 0.03878 0.0146 1.0000 0.9013
-4.250 -0.3236 0.04582 0.03941 0.0224 1.0000 0.8565
-4.000 -0.3602 0.04570 0.03953 0.0286 1.0000 0.8129
-3.750 -0.3947 0.04528 0.03934 0.0342 1.0000 0.7795
-3.500 -0.4314 0.04423 0.03856 0.0396 1.0000 0.7429
-3.250 -0.3004 0.03497 0.02633 -0.0344 1.0000 0.1804
-3.000 -0.2647 0.03200 0.02248 -0.0345 1.0000 0.1367
-2.750 -0.2345 0.02919 0.01906 -0.0336 1.0000 0.1155
-2.500 -0.2056 0.02690 0.01623 -0.0325 1.0000 0.1025
-2.250 -0.1766 0.02508 0.01379 -0.0311 1.0000 0.0946
-2.000 -0.1486 0.02321 0.01153 -0.0299 1.0000 0.0956
-1.750 -0.1203 0.02123 0.00932 -0.0287 1.0000 0.0992
-1.500 -0.0917 0.01962 0.00753 -0.0278 1.0000 0.1145
-1.250 -0.0651 0.01808 0.00599 -0.0270 1.0000 0.1540
-1.000 -0.0213 0.01448 0.00427 -0.0282 1.0000 1.0000
-0.750 0.0018 0.01451 0.00364 -0.0271 1.0000 1.0000
-0.500 0.0246 0.01456 0.00328 -0.0262 1.0000 1.0000
-0.250 0.0472 0.01463 0.00303 -0.0254 1.0000 1.0000
0.000 0.0698 0.01472 0.00288 -0.0245 1.0000 1.0000
0.250 0.0923 0.01482 0.00278 -0.0237 1.0000 1.0000
0.500 0.1147 0.01494 0.00279 -0.0230 1.0000 1.0000
0.750 0.1371 0.01508 0.00285 -0.0222 1.0000 1.0000
1.000 0.1592 0.01523 0.00300 -0.0214 1.0000 1.0000
1.250 0.1814 0.01540 0.00322 -0.0207 1.0000 1.0000
1.500 0.2033 0.01559 0.00350 -0.0199 1.0000 1.0000
1.750 0.2252 0.01580 0.00385 -0.0191 1.0000 1.0000
2.000 0.2469 0.01604 0.00429 -0.0183 1.0000 1.0000
2.250 0.2685 0.01630 0.00489 -0.0175 1.0000 1.0000
2.500 0.2900 0.01658 0.00549 -0.0168 1.0000 1.0000
2.750 0.3114 0.01691 0.00623 -0.0160 1.0000 1.0000
3.000 0.3325 0.01727 0.00721 -0.0150 1.0000 1.0000
3.250 0.3540 0.01766 0.00840 -0.0140 1.0000 1.0000
3.500 0.5369 0.02617 0.01402 -0.0332 0.0821 1.0000
3.750 0.5702 0.02878 0.01699 -0.0324 0.0809 1.0000
4.000 0.6008 0.03231 0.02075 -0.0317 0.0844 1.0000
4.250 0.6307 0.03462 0.02387 -0.0297 0.0955 1.0000
4.500 0.6591 0.03793 0.02782 -0.0280 0.1108 1.0000
4.750 0.6884 0.04129 0.03204 -0.0263 0.1376 1.0000
5.000 0.7224 0.04542 0.03727 -0.0257 0.1956 1.0000
5.250 0.7744 0.05366 0.04764 -0.0446 0.4131 1.0000
5.500 0.7559 0.06147 0.05574 -0.0633 0.5182 1.0000
5.750 0.7201 0.06643 0.06052 -0.0698 0.5761 1.0000
6.000 0.6762 0.06936 0.06317 -0.0717 0.6257 1.0000
7.000 0.7255 0.08528 0.07905 -0.0667 0.5181 1.0000
7.250 0.7525 0.08956 0.08336 -0.0615 0.4523 1.0000
7.500 0.7675 0.09309 0.08691 -0.0560 0.3915 1.0000
7.750 0.7787 0.09668 0.09049 -0.0520 0.3467 1.0000
8.000 0.7944 0.10103 0.09482 -0.0476 0.3037 1.0000
8.250 0.7828 0.10387 0.09757 -0.0479 0.2837 1.0000
8.500 0.7943 0.10847 0.10215 -0.0455 0.2585 1.0000
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Polar data table (+)
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