GOE 595 AIRFOIL (goe595-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: GOE 595 AIRFOIL (goe595-il) Reynolds number: 100,000 Max Cl/Cd: 55.1 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe595-il-100000.txt Download as CSV file: xf-goe595-il-100000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4130 0.10388 0.09899 -0.0361 1.0000 0.0921
-9.000 -0.3973 0.10041 0.09548 -0.0334 1.0000 0.0956
-8.750 -0.3977 0.09767 0.09280 -0.0333 1.0000 0.0991
-8.500 -0.4083 0.09532 0.09055 -0.0343 1.0000 0.1024
-8.250 -0.4341 0.09378 0.08917 -0.0356 1.0000 0.1038
-8.000 -0.4626 0.09200 0.08754 -0.0369 1.0000 0.1043
-7.750 -0.4523 0.08761 0.08317 -0.0329 1.0000 0.1066
-7.500 -0.4442 0.08505 0.08064 -0.0295 1.0000 0.1098
-7.250 -0.4533 0.08257 0.07822 -0.0285 1.0000 0.1128
-7.000 -0.4713 0.07987 0.07556 -0.0300 1.0000 0.1167
-6.750 -0.4967 0.07670 0.07227 -0.0342 1.0000 0.1191
-6.500 -0.4832 0.07330 0.06904 -0.0289 1.0000 0.1217
-6.250 -0.4809 0.07080 0.06656 -0.0268 1.0000 0.1257
-6.000 -0.4916 0.06699 0.06251 -0.0306 1.0000 0.1339
-5.750 -0.4835 0.06402 0.05966 -0.0273 1.0000 0.1366
-5.500 -0.4829 0.06102 0.05637 -0.0291 1.0000 0.1481
-5.250 -0.4736 0.05826 0.05380 -0.0255 1.0000 0.1529
-5.000 -0.4670 0.05526 0.05068 -0.0255 1.0000 0.1652
-4.750 -0.4583 0.05256 0.04788 -0.0249 1.0000 0.1791
-4.500 -0.4480 0.04995 0.04519 -0.0239 1.0000 0.1937
-4.250 -0.4066 0.03733 0.03060 -0.0287 1.0000 0.0919
-4.000 -0.3822 0.03294 0.02555 -0.0275 1.0000 0.0747
-3.750 -0.3612 0.03002 0.02227 -0.0264 1.0000 0.0717
-3.500 -0.3396 0.02808 0.01986 -0.0253 1.0000 0.0724
-3.250 -0.3177 0.02652 0.01788 -0.0240 1.0000 0.0741
-3.000 -0.2951 0.02498 0.01598 -0.0228 1.0000 0.0746
-2.750 -0.2727 0.02388 0.01456 -0.0216 1.0000 0.0762
-2.500 -0.2515 0.02260 0.01332 -0.0208 1.0000 0.0819
-2.250 -0.2265 0.02185 0.01245 -0.0203 0.9990 0.0877
-2.000 -0.1857 0.02104 0.01169 -0.0229 0.9927 0.0975
-1.250 -0.0390 0.01715 0.01090 -0.0348 0.9788 1.0000
-1.000 0.0074 0.01765 0.01105 -0.0389 0.9711 1.0000
-0.750 0.0454 0.01791 0.01105 -0.0414 0.9609 1.0000
-0.500 0.0852 0.01821 0.01114 -0.0442 0.9514 1.0000
-0.250 0.1329 0.01853 0.01127 -0.0484 0.9437 1.0000
0.000 0.1689 0.01868 0.01128 -0.0503 0.9323 1.0000
0.250 0.2068 0.01884 0.01133 -0.0525 0.9215 1.0000
0.500 0.2517 0.01896 0.01134 -0.0559 0.9127 1.0000
0.750 0.2940 0.01898 0.01131 -0.0587 0.9025 1.0000
1.000 0.3313 0.01898 0.01126 -0.0605 0.8905 1.0000
1.250 0.3720 0.01889 0.01115 -0.0628 0.8791 1.0000
1.500 0.4328 0.01838 0.01064 -0.0685 0.8721 1.0000
1.750 0.4773 0.01790 0.01018 -0.0709 0.8589 1.0000
2.000 0.5209 0.01730 0.00961 -0.0729 0.8451 1.0000
2.250 0.5628 0.01666 0.00901 -0.0744 0.8305 1.0000
2.500 0.6012 0.01613 0.00851 -0.0754 0.8155 1.0000
2.750 0.6328 0.01580 0.00823 -0.0752 0.7979 1.0000
3.000 0.6631 0.01550 0.00797 -0.0749 0.7783 1.0000
3.250 0.6975 0.01517 0.00764 -0.0752 0.7591 1.0000
3.500 0.7250 0.01507 0.00757 -0.0744 0.7359 1.0000
3.750 0.7557 0.01497 0.00745 -0.0742 0.7136 1.0000
4.000 0.7814 0.01505 0.00750 -0.0732 0.6883 1.0000
4.250 0.8057 0.01521 0.00764 -0.0721 0.6628 1.0000
4.500 0.8301 0.01542 0.00784 -0.0709 0.6379 1.0000
4.750 0.8545 0.01566 0.00802 -0.0698 0.6132 1.0000
5.000 0.8757 0.01593 0.00823 -0.0680 0.5844 1.0000
5.250 0.8943 0.01623 0.00845 -0.0658 0.5516 1.0000
5.500 0.9126 0.01660 0.00873 -0.0636 0.5185 1.0000
5.750 0.9313 0.01704 0.00909 -0.0616 0.4882 1.0000
6.000 0.9496 0.01752 0.00952 -0.0596 0.4592 1.0000
6.250 0.9675 0.01799 0.00998 -0.0576 0.4317 1.0000
6.500 0.9834 0.01840 0.01043 -0.0552 0.4023 1.0000
6.750 0.9973 0.01876 0.01081 -0.0525 0.3699 1.0000
7.000 1.0070 0.01909 0.01111 -0.0490 0.3204 1.0000
7.250 1.0104 0.02007 0.01164 -0.0448 0.2406 1.0000
7.500 1.0097 0.02210 0.01280 -0.0404 0.1233 1.0000
7.750 1.0136 0.02394 0.01428 -0.0367 0.0946 1.0000
8.000 1.0207 0.02541 0.01575 -0.0333 0.0840 1.0000
8.250 1.0261 0.02694 0.01731 -0.0298 0.0766 1.0000
8.500 1.0326 0.02853 0.01886 -0.0267 0.0701 1.0000
8.750 1.0490 0.03052 0.02083 -0.0249 0.0661 1.0000
9.000 1.0745 0.03252 0.02290 -0.0245 0.0615 1.0000
9.250 1.1169 0.03652 0.02675 -0.0273 0.0559 1.0000
9.500 1.1420 0.03893 0.02949 -0.0266 0.0546 1.0000
9.750 1.1629 0.04167 0.03264 -0.0254 0.0535 1.0000
10.000 1.1760 0.04424 0.03560 -0.0233 0.0521 1.0000
10.250 1.1847 0.04690 0.03862 -0.0208 0.0505 1.0000
10.500 1.1895 0.05008 0.04222 -0.0179 0.0506 1.0000
10.750 1.1894 0.05350 0.04605 -0.0147 0.0514 1.0000
11.000 1.1845 0.05711 0.05004 -0.0113 0.0523 1.0000
11.250 1.1763 0.06071 0.05393 -0.0079 0.0534 1.0000
11.500 1.1661 0.06466 0.05812 -0.0048 0.0544 1.0000
11.750 1.1759 0.06920 0.06286 -0.0037 0.0573 1.0000
12.000 1.1421 0.07119 0.06521 0.0009 0.0582 1.0000
12.250 1.1055 0.07484 0.06918 0.0032 0.0591 1.0000
12.500 1.0716 0.07971 0.07432 0.0033 0.0598 1.0000
12.750 1.0380 0.08573 0.08055 0.0013 0.0605 1.0000
13.000 1.0053 0.09288 0.08787 -0.0026 0.0609 1.0000
13.250 0.9678 0.10241 0.09752 -0.0092 0.0615 1.0000
13.500 0.9230 0.11639 0.11153 -0.0192 0.0633 1.0000
13.750 0.9077 0.12579 0.12089 -0.0237 0.0665 1.0000
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Polar data table (+)
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