GOE 595 AIRFOIL (goe595-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: GOE 595 AIRFOIL (goe595-il) Reynolds number: 100,000 Max Cl/Cd: 54.56 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe595-il-100000-n5.txt Download as CSV file: xf-goe595-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 595 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4041 0.08972 0.08499 -0.0386 1.0000 0.0336
-8.250 -0.4094 0.08701 0.08236 -0.0375 1.0000 0.0326
-8.000 -0.4216 0.08433 0.07978 -0.0365 1.0000 0.0324
-7.750 -0.4361 0.08201 0.07756 -0.0348 1.0000 0.0312
-7.250 -0.4605 0.07545 0.07112 -0.0345 1.0000 0.0303
-6.750 -0.4790 0.06223 0.05761 -0.0390 1.0000 0.0253
-6.500 -0.4685 0.05718 0.05241 -0.0413 0.9967 0.0249
-6.250 -0.4503 0.05154 0.04648 -0.0450 0.9915 0.0246
-6.000 -0.4291 0.04606 0.04060 -0.0482 0.9867 0.0243
-5.750 -0.4089 0.04091 0.03493 -0.0499 0.9812 0.0244
-5.500 -0.3836 0.03650 0.02987 -0.0515 0.9772 0.0245
-5.250 -0.3606 0.03314 0.02583 -0.0517 0.9718 0.0257
-5.000 -0.3345 0.03056 0.02297 -0.0529 0.9679 0.0279
-4.750 -0.3067 0.02825 0.02021 -0.0533 0.9641 0.0291
-4.500 -0.2797 0.02621 0.01774 -0.0533 0.9594 0.0303
-4.250 -0.2485 0.02459 0.01571 -0.0539 0.9558 0.0332
-4.000 -0.2158 0.02300 0.01382 -0.0549 0.9531 0.0362
-3.750 -0.1915 0.02182 0.01255 -0.0543 0.9472 0.0387
-3.500 -0.1608 0.02095 0.01156 -0.0549 0.9429 0.0440
-3.250 -0.1266 0.02000 0.01055 -0.0563 0.9395 0.0491
-3.000 -0.1004 0.01944 0.00989 -0.0560 0.9318 0.0560
-2.750 -0.0632 0.01877 0.00915 -0.0577 0.9265 0.0686
-2.500 -0.0333 0.01805 0.00848 -0.0580 0.9178 0.0901
-2.250 0.0045 0.01710 0.00791 -0.0601 0.9121 0.1780
-2.000 0.0320 0.01642 0.00759 -0.0602 0.9037 0.2922
-1.750 0.0654 0.01563 0.00734 -0.0615 0.8989 0.4507
-1.500 0.0936 0.01473 0.00730 -0.0609 0.8947 0.6919
-1.250 0.2017 0.01435 0.00725 -0.0764 0.9004 1.0000
-1.000 0.2362 0.01426 0.00696 -0.0777 0.8947 1.0000
-0.750 0.2626 0.01422 0.00678 -0.0774 0.8856 1.0000
-0.500 0.2924 0.01416 0.00658 -0.0778 0.8779 1.0000
-0.250 0.3227 0.01409 0.00638 -0.0782 0.8699 1.0000
0.000 0.3482 0.01407 0.00627 -0.0777 0.8600 1.0000
0.250 0.3816 0.01396 0.00607 -0.0786 0.8528 1.0000
0.500 0.4062 0.01395 0.00599 -0.0778 0.8417 1.0000
0.750 0.4313 0.01393 0.00592 -0.0771 0.8304 1.0000
1.000 0.4582 0.01388 0.00583 -0.0767 0.8187 1.0000
1.250 0.4856 0.01381 0.00571 -0.0764 0.8059 1.0000
1.500 0.5125 0.01372 0.00558 -0.0758 0.7902 1.0000
1.750 0.5403 0.01363 0.00546 -0.0754 0.7730 1.0000
2.000 0.5697 0.01353 0.00529 -0.0753 0.7544 1.0000
2.250 0.5960 0.01351 0.00522 -0.0746 0.7330 1.0000
2.500 0.6240 0.01349 0.00514 -0.0743 0.7103 1.0000
2.750 0.6492 0.01357 0.00516 -0.0735 0.6867 1.0000
3.000 0.6752 0.01367 0.00521 -0.0728 0.6645 1.0000
3.250 0.6995 0.01382 0.00531 -0.0719 0.6402 1.0000
3.500 0.7235 0.01400 0.00546 -0.0710 0.6154 1.0000
3.750 0.7475 0.01422 0.00561 -0.0701 0.5914 1.0000
4.000 0.7708 0.01447 0.00581 -0.0691 0.5683 1.0000
4.250 0.7938 0.01476 0.00607 -0.0680 0.5462 1.0000
4.500 0.8163 0.01508 0.00635 -0.0669 0.5246 1.0000
4.750 0.8388 0.01543 0.00666 -0.0658 0.5047 1.0000
5.000 0.8610 0.01578 0.00702 -0.0647 0.4858 1.0000
5.250 0.8818 0.01620 0.00741 -0.0633 0.4630 1.0000
5.500 0.9009 0.01664 0.00780 -0.0617 0.4355 1.0000
5.750 0.9191 0.01710 0.00821 -0.0599 0.4066 1.0000
6.000 0.9379 0.01754 0.00865 -0.0583 0.3798 1.0000
6.250 0.9571 0.01798 0.00915 -0.0568 0.3558 1.0000
6.500 0.9764 0.01842 0.00965 -0.0552 0.3321 1.0000
6.750 0.9938 0.01893 0.01015 -0.0535 0.3029 1.0000
7.000 1.0082 0.01962 0.01070 -0.0512 0.2659 1.0000
7.250 1.0223 0.02041 0.01138 -0.0491 0.2323 1.0000
7.500 1.0343 0.02141 0.01219 -0.0467 0.1781 1.0000
7.750 1.0363 0.02333 0.01341 -0.0432 0.0820 1.0000
8.000 1.0433 0.02490 0.01481 -0.0403 0.0625 1.0000
8.250 1.0494 0.02639 0.01632 -0.0371 0.0531 1.0000
8.500 1.0578 0.02755 0.01765 -0.0342 0.0462 1.0000
8.750 1.0598 0.02913 0.01931 -0.0307 0.0413 1.0000
9.000 1.0650 0.03051 0.02087 -0.0276 0.0375 1.0000
9.250 1.0704 0.03193 0.02240 -0.0248 0.0338 1.0000
9.500 1.0723 0.03365 0.02417 -0.0220 0.0312 1.0000
9.750 1.0767 0.03546 0.02609 -0.0193 0.0294 1.0000
10.000 1.0857 0.03706 0.02787 -0.0173 0.0271 1.0000
10.250 1.0940 0.03871 0.02966 -0.0154 0.0247 1.0000
10.500 1.1017 0.04050 0.03158 -0.0137 0.0230 1.0000
10.750 1.1114 0.04271 0.03383 -0.0122 0.0218 1.0000
11.000 1.1267 0.04534 0.03661 -0.0110 0.0208 1.0000
11.250 1.1361 0.04757 0.03915 -0.0095 0.0198 1.0000
11.500 1.1427 0.05008 0.04194 -0.0079 0.0188 1.0000
11.750 1.1451 0.05268 0.04483 -0.0063 0.0179 1.0000
12.000 1.1449 0.05564 0.04806 -0.0048 0.0173 1.0000
12.250 1.1413 0.05893 0.05163 -0.0035 0.0170 1.0000
12.500 1.1345 0.06244 0.05541 -0.0025 0.0167 1.0000
12.750 1.1248 0.06631 0.05955 -0.0019 0.0166 1.0000
13.000 1.1120 0.07063 0.06412 -0.0019 0.0164 1.0000
13.250 1.0973 0.07539 0.06913 -0.0026 0.0163 1.0000
13.500 1.0813 0.08061 0.07459 -0.0041 0.0164 1.0000
13.750 1.0633 0.08648 0.08070 -0.0064 0.0164 1.0000
14.000 1.0435 0.09311 0.08754 -0.0096 0.0165 1.0000
14.250 1.0235 0.10041 0.09504 -0.0137 0.0166 1.0000
14.500 1.0016 0.10878 0.10360 -0.0188 0.0168 1.0000
14.750 0.9788 0.11819 0.11316 -0.0249 0.0172 1.0000
15.000 0.9542 0.12915 0.12423 -0.0319 0.0175 1.0000
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Polar data table (+)
Polar graphs
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