GOE 15K AIRFOIL (goe15k-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 15K AIRFOIL (goe15k-il) Reynolds number: 100,000 Max Cl/Cd: 30.73 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe15k-il-100000-n5.txt Download as CSV file: xf-goe15k-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 15K 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.4386 0.11014 0.10483 -0.0830 0.9580 0.0434
-11.750 -0.4626 0.10088 0.09561 -0.0872 0.9559 0.0433
-11.500 -0.5286 0.08431 0.07910 -0.0948 0.9514 0.0428
-11.250 -0.5905 0.07181 0.06641 -0.1026 0.9473 0.0411
-11.000 -0.6325 0.06630 0.06066 -0.1012 0.9439 0.0409
-10.750 -0.6689 0.06329 0.05748 -0.0953 0.9387 0.0405
-10.500 -0.6965 0.06042 0.05435 -0.0900 0.9349 0.0412
-10.250 -0.7202 0.05767 0.05131 -0.0842 0.9323 0.0414
-10.000 -0.7537 0.05664 0.05015 -0.0744 0.9270 0.0416
-9.750 -0.7739 0.05478 0.04803 -0.0673 0.9235 0.0422
-9.500 -0.7879 0.05227 0.04508 -0.0613 0.9206 0.0434
-9.250 -0.8026 0.05005 0.04234 -0.0545 0.9174 0.0449
-9.000 -0.8149 0.04880 0.04090 -0.0477 0.9139 0.0458
-8.750 -0.8127 0.04768 0.03971 -0.0437 0.9115 0.0469
-8.500 -0.8066 0.04684 0.03878 -0.0403 0.9092 0.0486
-8.250 -0.7981 0.04560 0.03731 -0.0372 0.9070 0.0505
-8.000 -0.7886 0.04425 0.03557 -0.0342 0.9053 0.0535
-7.750 -0.7871 0.04317 0.03406 -0.0293 0.9030 0.0555
-7.500 -0.7865 0.04201 0.03271 -0.0244 0.9002 0.0570
-7.250 -0.7761 0.04105 0.03174 -0.0216 0.8977 0.0592
-7.000 -0.7633 0.04042 0.03098 -0.0191 0.8952 0.0621
-6.750 -0.7474 0.03954 0.02984 -0.0170 0.8934 0.0650
-6.500 -0.7300 0.03885 0.02879 -0.0151 0.8919 0.0680
-6.250 -0.7062 0.03774 0.02765 -0.0150 0.8908 0.0712
-6.000 -0.6810 0.03709 0.02692 -0.0149 0.8896 0.0747
-5.750 -0.6749 0.03654 0.02625 -0.0108 0.8860 0.0768
-5.500 -0.6608 0.03609 0.02561 -0.0084 0.8834 0.0798
-5.250 -0.6389 0.03545 0.02480 -0.0075 0.8812 0.0822
-5.000 -0.6148 0.03475 0.02412 -0.0072 0.8793 0.0852
-4.750 -0.5894 0.03425 0.02356 -0.0071 0.8775 0.0878
-4.500 -0.5627 0.03383 0.02305 -0.0072 0.8760 0.0918
-4.250 -0.5340 0.03352 0.02259 -0.0076 0.8747 0.0955
-4.000 -0.5220 0.03311 0.02222 -0.0048 0.8713 0.0980
-3.750 -0.5089 0.03282 0.02189 -0.0022 0.8676 0.1011
-3.500 -0.4877 0.03255 0.02156 -0.0012 0.8648 0.1044
-3.250 -0.4629 0.03231 0.02127 -0.0009 0.8627 0.1109
-3.000 -0.4351 0.03203 0.02103 -0.0012 0.8609 0.1197
-2.750 -0.4042 0.03178 0.02083 -0.0022 0.8594 0.1337
-2.500 -0.3697 0.03147 0.02069 -0.0040 0.8583 0.1679
-2.250 -0.3627 0.03131 0.02069 -0.0005 0.8531 0.2011
-2.000 -0.3357 0.03113 0.02072 -0.0011 0.8503 0.2549
-1.750 -0.3013 0.03101 0.02079 -0.0031 0.8482 0.3086
-1.500 -0.2613 0.03096 0.02089 -0.0064 0.8466 0.3601
-1.250 0.0131 0.03284 0.02494 -0.0594 0.8604 0.9127
-1.000 0.0459 0.03400 0.02598 -0.0603 0.8593 0.9621
-0.750 0.1042 0.03497 0.02679 -0.0672 0.8590 0.9817
-0.500 0.1692 0.03554 0.02725 -0.0758 0.8589 0.9951
-0.250 0.2024 0.03588 0.02749 -0.0779 0.8564 1.0000
0.000 0.2078 0.03620 0.02776 -0.0742 0.8508 1.0000
0.250 0.2290 0.03631 0.02780 -0.0734 0.8474 1.0000
0.500 0.2544 0.03638 0.02780 -0.0734 0.8450 1.0000
0.750 0.2822 0.03645 0.02781 -0.0738 0.8432 1.0000
1.000 0.2845 0.03692 0.02826 -0.0695 0.8368 1.0000
1.250 0.3027 0.03712 0.02842 -0.0682 0.8326 1.0000
1.500 0.3283 0.03720 0.02847 -0.0682 0.8297 1.0000
1.750 0.3576 0.03723 0.02847 -0.0687 0.8276 1.0000
2.000 0.3631 0.03760 0.02883 -0.0650 0.8196 1.0000
2.250 0.3880 0.03759 0.02882 -0.0647 0.8151 1.0000
2.500 0.4220 0.03733 0.02855 -0.0659 0.8120 1.0000
2.750 0.4300 0.03768 0.02892 -0.0626 0.8039 1.0000
3.000 0.4547 0.03767 0.02892 -0.0623 0.7992 1.0000
3.250 0.4872 0.03743 0.02871 -0.0632 0.7963 1.0000
3.500 0.4963 0.03769 0.02900 -0.0600 0.7868 1.0000
3.750 0.5252 0.03747 0.02882 -0.0602 0.7824 1.0000
4.000 0.5585 0.03718 0.02857 -0.0612 0.7798 1.0000
4.250 0.5619 0.03770 0.02915 -0.0572 0.7692 1.0000
4.500 0.5917 0.03749 0.02901 -0.0576 0.7658 1.0000
5.000 0.6322 0.03736 0.02902 -0.0549 0.7509 1.0000
5.250 0.6467 0.03729 0.02902 -0.0524 0.7394 1.0000
5.500 0.6711 0.03655 0.02835 -0.0511 0.7275 1.0000
6.500 0.7473 0.03516 0.02735 -0.0434 0.6758 1.0000
7.000 0.7848 0.03425 0.02657 -0.0389 0.6307 1.0000
7.250 0.8179 0.03292 0.02523 -0.0384 0.5977 1.0000
7.500 0.8661 0.03083 0.02299 -0.0399 0.5379 1.0000
7.750 0.9074 0.02953 0.02103 -0.0407 0.4451 1.0000
8.250 0.9309 0.03082 0.02186 -0.0355 0.3518 1.0000
8.500 0.9385 0.03171 0.02256 -0.0325 0.3052 1.0000
8.750 0.9443 0.03275 0.02334 -0.0294 0.2601 1.0000
9.000 0.9505 0.03385 0.02418 -0.0264 0.2223 1.0000
9.250 0.9593 0.03482 0.02502 -0.0239 0.1963 1.0000
9.500 0.9686 0.03581 0.02591 -0.0214 0.1746 1.0000
9.750 0.9774 0.03685 0.02687 -0.0190 0.1564 1.0000
10.000 0.9871 0.03784 0.02782 -0.0167 0.1407 1.0000
10.250 0.9973 0.03881 0.02878 -0.0146 0.1265 1.0000
10.500 1.0080 0.03976 0.02975 -0.0126 0.1151 1.0000
10.750 1.0172 0.04083 0.03082 -0.0104 0.1049 1.0000
11.000 1.0256 0.04196 0.03194 -0.0082 0.0967 1.0000
11.250 1.0357 0.04300 0.03306 -0.0062 0.0887 1.0000
11.500 1.0429 0.04427 0.03431 -0.0040 0.0809 1.0000
11.750 1.0517 0.04544 0.03558 -0.0020 0.0742 1.0000
12.000 1.0585 0.04680 0.03695 0.0002 0.0693 1.0000
12.250 1.0666 0.04808 0.03834 0.0021 0.0630 1.0000
12.500 1.0729 0.04956 0.03985 0.0042 0.0590 1.0000
12.750 1.0806 0.05096 0.04138 0.0061 0.0541 1.0000
13.000 1.0845 0.05264 0.04305 0.0082 0.0506 1.0000
13.250 1.0934 0.05408 0.04467 0.0100 0.0476 1.0000
13.500 1.0997 0.05567 0.04638 0.0118 0.0442 1.0000
13.750 1.1030 0.05745 0.04822 0.0136 0.0417 1.0000
14.000 1.1080 0.05926 0.05011 0.0153 0.0400 1.0000
14.250 1.1150 0.06100 0.05208 0.0169 0.0385 1.0000
14.500 1.1182 0.06298 0.05424 0.0186 0.0361 1.0000
14.750 1.1215 0.06500 0.05639 0.0202 0.0350 1.0000
15.000 1.1207 0.06733 0.05879 0.0217 0.0333 1.0000
15.250 1.1220 0.06963 0.06119 0.0231 0.0326 1.0000
15.500 1.1226 0.07221 0.06388 0.0245 0.0317 1.0000
15.750 1.1209 0.07504 0.06700 0.0258 0.0309 1.0000
16.000 1.1157 0.07829 0.07050 0.0270 0.0298 1.0000
16.250 1.1093 0.08171 0.07415 0.0279 0.0290 1.0000
16.500 1.1026 0.08530 0.07795 0.0286 0.0287 1.0000
16.750 1.0929 0.08913 0.08194 0.0288 0.0276 1.0000
17.000 1.0827 0.09336 0.08635 0.0288 0.0271 1.0000
17.250 1.0701 0.09816 0.09139 0.0285 0.0272 1.0000
17.500 1.0570 0.10309 0.09649 0.0276 0.0267 1.0000
17.750 1.0417 0.10868 0.10227 0.0261 0.0266 1.0000
18.000 1.0185 0.11606 0.10992 0.0235 0.0267 1.0000
18.250 1.0043 0.12196 0.11592 0.0210 0.0261 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 15K AIRFOIL (goe15k-il)