FX 67-K-170/17 AIRFOIL (fx67k170-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 67-K-170/17 AIRFOIL (fx67k170-il) Reynolds number: 500,000 Max Cl/Cd: 111.39 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx67k170-il-500000-n5.txt Download as CSV file: xf-fx67k170-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 67-K-170/17 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.1124 0.08455 0.08089 -0.1075 0.7203 0.0069
-10.750 -0.1176 0.07967 0.07604 -0.1098 0.7188 0.0068
-10.500 -0.1269 0.07410 0.07050 -0.1124 0.7172 0.0067
-10.250 -0.1393 0.06794 0.06438 -0.1158 0.7158 0.0066
-10.000 -0.1773 0.05592 0.05235 -0.1248 0.7143 0.0064
-9.750 -0.2171 0.04884 0.04513 -0.1268 0.7126 0.0063
-9.500 -0.2535 0.04380 0.03994 -0.1245 0.7107 0.0062
-9.250 -0.3031 0.03681 0.03260 -0.1188 0.7086 0.0061
-9.000 -0.3470 0.02764 0.02261 -0.1123 0.7067 0.0064
-8.750 -0.3429 0.02439 0.01888 -0.1100 0.7049 0.0068
-8.500 -0.3303 0.02209 0.01620 -0.1084 0.7032 0.0070
-8.250 -0.3129 0.02043 0.01424 -0.1073 0.7016 0.0072
-8.000 -0.2924 0.01928 0.01285 -0.1066 0.7001 0.0074
-7.750 -0.2716 0.01818 0.01155 -0.1059 0.6987 0.0077
-7.500 -0.2482 0.01764 0.01095 -0.1056 0.6972 0.0080
-7.250 -0.2237 0.01722 0.01049 -0.1055 0.6956 0.0083
-7.000 -0.1985 0.01693 0.01015 -0.1054 0.6939 0.0088
-6.750 -0.1742 0.01637 0.00950 -0.1051 0.6921 0.0093
-6.500 -0.1505 0.01569 0.00870 -0.1047 0.6903 0.0098
-6.250 -0.1266 0.01507 0.00797 -0.1043 0.6886 0.0102
-6.000 -0.1022 0.01457 0.00741 -0.1041 0.6871 0.0108
-5.750 -0.0764 0.01428 0.00708 -0.1040 0.6857 0.0114
-5.500 -0.0509 0.01394 0.00668 -0.1039 0.6844 0.0122
-5.250 -0.0254 0.01360 0.00629 -0.1038 0.6828 0.0134
-5.000 -0.0002 0.01323 0.00590 -0.1037 0.6810 0.0145
-4.750 0.0259 0.01298 0.00564 -0.1037 0.6792 0.0159
-4.500 0.0526 0.01278 0.00540 -0.1038 0.6773 0.0179
-4.250 0.0784 0.01246 0.00503 -0.1037 0.6755 0.0196
-4.000 0.1045 0.01217 0.00472 -0.1037 0.6740 0.0211
-3.750 0.1313 0.01196 0.00446 -0.1038 0.6725 0.0228
-3.500 0.1585 0.01181 0.00425 -0.1040 0.6712 0.0246
-3.250 0.1850 0.01157 0.00399 -0.1041 0.6697 0.0269
-3.000 0.2120 0.01141 0.00382 -0.1042 0.6680 0.0287
-2.750 0.2392 0.01127 0.00367 -0.1044 0.6661 0.0312
-2.500 0.2664 0.01113 0.00352 -0.1045 0.6642 0.0357
-2.250 0.2933 0.01092 0.00338 -0.1047 0.6623 0.0570
-2.000 0.3186 0.01051 0.00322 -0.1047 0.6607 0.1376
-1.750 0.3429 0.00995 0.00307 -0.1047 0.6592 0.2688
-1.500 0.3649 0.00916 0.00292 -0.1045 0.6577 0.4678
-1.250 0.3868 0.00856 0.00291 -0.1038 0.6563 0.6416
-1.000 0.4112 0.00849 0.00311 -0.1032 0.6545 0.7255
-0.750 0.4386 0.00855 0.00320 -0.1032 0.6525 0.7505
-0.500 0.4664 0.00862 0.00326 -0.1034 0.6505 0.7651
-0.250 0.4944 0.00869 0.00329 -0.1036 0.6487 0.7769
0.250 0.5497 0.00884 0.00343 -0.1038 0.6455 0.7979
0.500 0.5780 0.00890 0.00344 -0.1042 0.6440 0.8034
0.750 0.6067 0.00896 0.00344 -0.1047 0.6425 0.8082
1.000 0.6340 0.00902 0.00353 -0.1048 0.6405 0.8114
1.250 0.6615 0.00908 0.00360 -0.1051 0.6384 0.8151
1.500 0.6893 0.00914 0.00366 -0.1054 0.6365 0.8191
1.750 0.7173 0.00921 0.00371 -0.1058 0.6347 0.8229
2.000 0.7449 0.00925 0.00377 -0.1060 0.6329 0.8256
2.250 0.7727 0.00930 0.00381 -0.1063 0.6311 0.8283
2.500 0.8009 0.00936 0.00385 -0.1067 0.6294 0.8312
2.750 0.8287 0.00943 0.00393 -0.1071 0.6277 0.8340
3.000 0.8550 0.00950 0.00403 -0.1072 0.6241 0.8369
3.250 0.8813 0.00952 0.00407 -0.1073 0.6196 0.8392
3.500 0.9083 0.00954 0.00407 -0.1074 0.6157 0.8415
3.750 0.9347 0.00961 0.00418 -0.1075 0.6126 0.8442
4.000 0.9607 0.00970 0.00431 -0.1076 0.6099 0.8471
4.250 0.9868 0.00978 0.00442 -0.1076 0.6064 0.8499
4.500 1.0133 0.00985 0.00451 -0.1078 0.6031 0.8524
4.750 1.0392 0.00991 0.00458 -0.1078 0.5997 0.8545
5.000 1.0627 0.01000 0.00474 -0.1073 0.5940 0.8570
5.250 1.0868 0.01008 0.00485 -0.1070 0.5886 0.8595
5.500 1.1111 0.01019 0.00498 -0.1068 0.5843 0.8622
5.750 1.1340 0.01031 0.00517 -0.1063 0.5790 0.8651
6.000 1.1567 0.01044 0.00531 -0.1058 0.5736 0.8677
6.250 1.1774 0.01057 0.00552 -0.1048 0.5674 0.8702
6.500 1.1960 0.01074 0.00571 -0.1035 0.5600 0.8735
6.750 1.2119 0.01094 0.00596 -0.1017 0.5510 0.8774
7.000 1.2198 0.01121 0.00623 -0.0984 0.5397 0.8819
7.250 1.2241 0.01162 0.00668 -0.0946 0.5278 0.8870
7.500 1.2228 0.01232 0.00738 -0.0902 0.5112 0.8947
7.750 1.2157 0.01339 0.00842 -0.0853 0.4889 0.9054
8.000 1.2082 0.01472 0.00971 -0.0808 0.4633 0.9218
8.250 1.2081 0.01701 0.01186 -0.0797 0.4206 0.9643
8.500 1.1976 0.01885 0.01357 -0.0755 0.3939 1.0000
8.750 1.1824 0.02116 0.01570 -0.0711 0.3605 1.0000
9.000 1.1762 0.02307 0.01749 -0.0678 0.3347 1.0000
9.250 1.1683 0.02519 0.01946 -0.0646 0.3047 1.0000
9.500 1.1634 0.02728 0.02141 -0.0619 0.2755 1.0000
9.750 1.1593 0.02942 0.02340 -0.0594 0.2454 1.0000
10.000 1.1579 0.03142 0.02527 -0.0572 0.2173 1.0000
10.250 1.1499 0.03397 0.02759 -0.0546 0.1770 1.0000
10.500 1.1388 0.03684 0.03015 -0.0519 0.1298 1.0000
10.750 1.1383 0.03898 0.03213 -0.0502 0.0995 1.0000
11.000 1.1401 0.04102 0.03404 -0.0488 0.0736 1.0000
11.250 1.1390 0.04334 0.03619 -0.0472 0.0465 1.0000
11.500 1.1484 0.04485 0.03769 -0.0464 0.0409 1.0000
11.750 1.1583 0.04633 0.03918 -0.0457 0.0368 1.0000
12.000 1.1696 0.04774 0.04063 -0.0452 0.0345 1.0000
12.250 1.1813 0.04910 0.04205 -0.0447 0.0335 1.0000
12.500 1.1922 0.05055 0.04354 -0.0442 0.0319 1.0000
12.750 1.2013 0.05220 0.04522 -0.0436 0.0298 1.0000
13.000 1.2099 0.05390 0.04696 -0.0430 0.0272 1.0000
13.250 1.2230 0.05520 0.04833 -0.0427 0.0256 1.0000
13.500 1.2336 0.05676 0.04994 -0.0424 0.0215 1.0000
13.750 1.2388 0.05885 0.05194 -0.0418 0.0119 1.0000
14.000 1.2417 0.06121 0.05427 -0.0412 0.0066 1.0000
14.250 1.2470 0.06339 0.05650 -0.0407 0.0050 1.0000
14.500 1.2527 0.06555 0.05870 -0.0403 0.0042 1.0000
14.750 1.2584 0.06777 0.06099 -0.0399 0.0037 1.0000
15.000 1.2656 0.06984 0.06314 -0.0397 0.0034 1.0000
15.250 1.2715 0.07207 0.06545 -0.0395 0.0032 1.0000
15.500 1.2766 0.07444 0.06790 -0.0393 0.0031 1.0000
15.750 1.2827 0.07672 0.07027 -0.0392 0.0030 1.0000
16.000 1.2872 0.07923 0.07286 -0.0391 0.0029 1.0000
16.250 1.2908 0.08186 0.07557 -0.0391 0.0027 1.0000
16.500 1.2928 0.08475 0.07855 -0.0392 0.0026 1.0000
16.750 1.2969 0.08739 0.08129 -0.0393 0.0025 1.0000
17.000 1.3001 0.09021 0.08420 -0.0396 0.0025 1.0000
17.250 1.3022 0.09321 0.08731 -0.0399 0.0024 1.0000
17.500 1.3037 0.09632 0.09052 -0.0403 0.0023 1.0000
17.750 1.3054 0.09945 0.09375 -0.0408 0.0023 1.0000
18.000 1.3058 0.10280 0.09720 -0.0414 0.0022 1.0000
18.250 1.3060 0.10621 0.10073 -0.0421 0.0023 1.0000
18.500 1.3056 0.10974 0.10436 -0.0429 0.0022 1.0000
18.750 1.3044 0.11346 0.10818 -0.0438 0.0021 1.0000
19.000 1.3033 0.11719 0.11202 -0.0449 0.0021 1.0000
19.250 1.3021 0.12096 0.11589 -0.0461 0.0020 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 67-K-170/17 AIRFOIL (fx67k170-il)