FX 67-K-150/17 AIRFOIL (fx67k150-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 67-K-150/17 AIRFOIL (fx67k150-il) Reynolds number: 500,000 Max Cl/Cd: 118.25 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx67k150-il-500000-n5.txt Download as CSV file: xf-fx67k150-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 67-K-150/17 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.0977 0.07743 0.07391 -0.1012 0.7223 0.0046
-9.000 -0.0975 0.07393 0.07043 -0.1029 0.7205 0.0044
-8.750 -0.0975 0.07040 0.06693 -0.1048 0.7187 0.0044
-8.500 -0.1018 0.06618 0.06273 -0.1072 0.7169 0.0042
-8.250 -0.1101 0.06131 0.05788 -0.1112 0.7152 0.0042
-8.000 -0.1249 0.05704 0.05360 -0.1135 0.7133 0.0041
-7.750 -0.1356 0.05373 0.05024 -0.1136 0.7111 0.0041
-7.500 -0.1416 0.04982 0.04624 -0.1138 0.7090 0.0040
-7.250 -0.1420 0.04616 0.04245 -0.1136 0.7069 0.0038
-7.000 -0.1394 0.04238 0.03850 -0.1130 0.7050 0.0037
-6.750 -0.1337 0.03856 0.03446 -0.1121 0.7034 0.0036
-6.500 -0.1252 0.03465 0.03024 -0.1109 0.7019 0.0034
-6.250 -0.1141 0.03067 0.02593 -0.1095 0.7004 0.0032
-6.000 -0.1024 0.02566 0.02047 -0.1073 0.6986 0.0030
-5.750 -0.0891 0.01976 0.01386 -0.1046 0.6967 0.0029
-5.500 -0.0671 0.01723 0.01091 -0.1036 0.6948 0.0029
-5.250 -0.0428 0.01586 0.00929 -0.1031 0.6930 0.0031
-5.000 -0.0180 0.01479 0.00803 -0.1027 0.6913 0.0033
-4.750 0.0072 0.01397 0.00706 -0.1024 0.6898 0.0036
-4.500 0.0325 0.01327 0.00622 -0.1021 0.6884 0.0041
-4.250 0.0583 0.01275 0.00559 -0.1020 0.6869 0.0046
-4.000 0.0837 0.01217 0.00495 -0.1018 0.6850 0.0054
-3.750 0.1108 0.01195 0.00475 -0.1020 0.6829 0.0068
-3.500 0.1378 0.01168 0.00442 -0.1021 0.6809 0.0089
-3.250 0.1649 0.01143 0.00417 -0.1023 0.6791 0.0119
-3.000 0.1920 0.01115 0.00383 -0.1024 0.6776 0.0165
-2.750 0.2195 0.01093 0.00356 -0.1026 0.6761 0.0208
-2.500 0.2473 0.01080 0.00336 -0.1028 0.6746 0.0236
-2.250 0.2747 0.01059 0.00311 -0.1030 0.6728 0.0297
-2.000 0.3022 0.01040 0.00294 -0.1032 0.6707 0.0446
-1.750 0.3280 0.00988 0.00278 -0.1034 0.6686 0.1570
-1.500 0.3523 0.00913 0.00263 -0.1036 0.6667 0.3557
-1.250 0.3743 0.00822 0.00256 -0.1033 0.6650 0.6115
-1.000 0.3987 0.00800 0.00268 -0.1026 0.6633 0.7263
-0.750 0.4253 0.00806 0.00278 -0.1023 0.6617 0.7753
-0.500 0.4527 0.00812 0.00283 -0.1024 0.6598 0.7915
-0.250 0.4808 0.00817 0.00286 -0.1026 0.6578 0.7996
0.000 0.5088 0.00821 0.00287 -0.1029 0.6557 0.8059
0.250 0.5369 0.00825 0.00289 -0.1032 0.6536 0.8121
0.500 0.5652 0.00829 0.00289 -0.1035 0.6516 0.8178
0.750 0.5932 0.00832 0.00290 -0.1038 0.6499 0.8226
1.000 0.6216 0.00837 0.00291 -0.1041 0.6483 0.8277
1.250 0.6495 0.00842 0.00296 -0.1045 0.6462 0.8321
1.500 0.6771 0.00847 0.00303 -0.1047 0.6439 0.8361
1.750 0.7051 0.00852 0.00309 -0.1050 0.6416 0.8402
2.000 0.7333 0.00857 0.00314 -0.1054 0.6395 0.8441
2.250 0.7610 0.00860 0.00318 -0.1057 0.6374 0.8472
2.500 0.7889 0.00865 0.00322 -0.1060 0.6356 0.8506
2.750 0.8167 0.00871 0.00331 -0.1063 0.6335 0.8543
3.000 0.8443 0.00878 0.00341 -0.1067 0.6308 0.8581
3.250 0.8713 0.00882 0.00351 -0.1068 0.6280 0.8611
3.500 0.8985 0.00886 0.00357 -0.1070 0.6247 0.8645
3.750 0.9256 0.00889 0.00361 -0.1072 0.6201 0.8680
4.000 0.9519 0.00893 0.00370 -0.1072 0.6131 0.8718
4.250 0.9783 0.00897 0.00374 -0.1072 0.6085 0.8750
4.500 1.0044 0.00904 0.00389 -0.1073 0.6042 0.8786
4.750 1.0308 0.00911 0.00402 -0.1074 0.5998 0.8824
5.000 1.0570 0.00919 0.00413 -0.1074 0.5951 0.8861
5.250 1.0822 0.00926 0.00430 -0.1072 0.5893 0.8897
5.500 1.1056 0.00935 0.00441 -0.1067 0.5778 0.8941
6.000 1.1299 0.01011 0.00489 -0.1015 0.4988 0.9077
6.250 1.1126 0.01125 0.00571 -0.0937 0.4315 0.9257
6.500 1.0894 0.01323 0.00740 -0.0867 0.3626 1.0000
7.000 1.0629 0.01733 0.01101 -0.0775 0.2661 1.0000
7.250 1.0533 0.01937 0.01280 -0.0736 0.2168 1.0000
7.500 1.0365 0.02192 0.01497 -0.0691 0.1503 1.0000
7.750 1.0054 0.02550 0.01794 -0.0634 0.0438 1.0000
8.000 1.0125 0.02682 0.01919 -0.0618 0.0261 1.0000
8.250 1.0173 0.02832 0.02062 -0.0599 0.0052 1.0000
8.500 1.0289 0.02941 0.02175 -0.0587 0.0034 1.0000
8.750 1.0420 0.03043 0.02283 -0.0578 0.0032 1.0000
9.000 1.0544 0.03153 0.02399 -0.0568 0.0031 1.0000
9.250 1.0663 0.03269 0.02523 -0.0558 0.0029 1.0000
9.500 1.0776 0.03393 0.02654 -0.0548 0.0028 1.0000
9.750 1.0886 0.03521 0.02789 -0.0538 0.0028 1.0000
10.000 1.0984 0.03660 0.02940 -0.0527 0.0026 1.0000
10.250 1.1082 0.03804 0.03089 -0.0518 0.0024 1.0000
10.500 1.1154 0.03971 0.03263 -0.0507 0.0020 1.0000
10.750 1.1209 0.04157 0.03458 -0.0495 0.0019 1.0000
11.000 1.1259 0.04353 0.03662 -0.0484 0.0017 1.0000
11.250 1.1345 0.04521 0.03837 -0.0476 0.0016 1.0000
11.500 1.1395 0.04726 0.04051 -0.0465 0.0016 1.0000
11.750 1.1483 0.04898 0.04230 -0.0459 0.0015 1.0000
12.000 1.1538 0.05104 0.04445 -0.0450 0.0014 1.0000
12.250 1.1589 0.05320 0.04670 -0.0442 0.0014 1.0000
12.500 1.1642 0.05537 0.04895 -0.0434 0.0013 1.0000
12.750 1.1693 0.05758 0.05126 -0.0427 0.0012 1.0000
13.000 1.1750 0.05978 0.05354 -0.0420 0.0012 1.0000
13.250 1.1809 0.06198 0.05584 -0.0414 0.0012 1.0000
13.500 1.1866 0.06424 0.05824 -0.0408 0.0011 1.0000
13.750 1.1926 0.06649 0.06058 -0.0402 0.0010 1.0000
14.000 1.1991 0.06872 0.06293 -0.0398 0.0010 1.0000
14.250 1.2050 0.07109 0.06541 -0.0393 0.0009 1.0000
14.500 1.2108 0.07352 0.06798 -0.0389 0.0009 1.0000
14.750 1.2162 0.07606 0.07065 -0.0386 0.0009 1.0000
15.000 1.2202 0.07883 0.07358 -0.0383 0.0009 1.0000
15.250 1.2227 0.08190 0.07682 -0.0381 0.0009 1.0000
15.500 1.2250 0.08508 0.08017 -0.0381 0.0009 1.0000
15.750 1.2242 0.08875 0.08402 -0.0383 0.0009 1.0000
16.000 1.2216 0.09278 0.08824 -0.0387 0.0009 1.0000
16.250 1.2171 0.09720 0.09287 -0.0392 0.0009 1.0000
16.500 1.2107 0.10201 0.09788 -0.0402 0.0009 1.0000
16.750 1.2019 0.10735 0.10343 -0.0415 0.0009 1.0000
17.000 1.1918 0.11306 0.10936 -0.0432 0.0009 1.0000
17.250 1.1795 0.11934 0.11586 -0.0453 0.0009 1.0000
17.500 1.1656 0.12615 0.12289 -0.0480 0.0009 1.0000
17.750 1.1510 0.13335 0.13029 -0.0512 0.0009 1.0000
18.000 1.1356 0.14105 0.13819 -0.0550 0.0009 1.0000
18.250 1.1205 0.14898 0.14631 -0.0592 0.0009 1.0000
18.500 1.1042 0.15764 0.15515 -0.0642 0.0010 1.0000
18.750 1.0877 0.16691 0.16458 -0.0697 0.0010 1.0000
19.000 1.0727 0.17644 0.17425 -0.0755 0.0010 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 67-K-150/17 AIRFOIL (fx67k150-il)