GOE 546 AIRFOIL (goe546-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 546 AIRFOIL (goe546-il) Reynolds number: 200,000 Max Cl/Cd: 76.21 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe546-il-200000-n5.txt Download as CSV file: xf-goe546-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 546 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4137 0.08588 0.08226 -0.0446 1.0000 0.0224
-9.250 -0.4283 0.08109 0.07754 -0.0455 1.0000 0.0227
-9.000 -0.4507 0.07642 0.07297 -0.0456 1.0000 0.0229
-8.750 -0.4753 0.06909 0.06575 -0.0503 0.9959 0.0232
-8.500 -0.5060 0.03930 0.03493 -0.0832 0.9745 0.0239
-8.250 -0.4813 0.03808 0.03363 -0.0848 0.9678 0.0246
-8.000 -0.4544 0.03558 0.03088 -0.0877 0.9635 0.0255
-7.750 -0.4362 0.03212 0.02699 -0.0890 0.9555 0.0267
-7.500 -0.4119 0.02798 0.02208 -0.0912 0.9515 0.0290
-7.250 -0.3900 0.02587 0.01961 -0.0914 0.9456 0.0303
-7.000 -0.3616 0.02467 0.01827 -0.0924 0.9417 0.0314
-6.750 -0.3302 0.02360 0.01701 -0.0938 0.9389 0.0329
-6.500 -0.2992 0.02235 0.01548 -0.0950 0.9359 0.0350
-6.250 -0.2736 0.02114 0.01394 -0.0949 0.9303 0.0363
-6.000 -0.2435 0.01983 0.01240 -0.0957 0.9265 0.0376
-5.750 -0.2113 0.01894 0.01147 -0.0970 0.9235 0.0394
-5.500 -0.1812 0.01818 0.01060 -0.0977 0.9196 0.0408
-5.250 -0.1532 0.01742 0.00972 -0.0978 0.9143 0.0421
-5.000 -0.1217 0.01673 0.00891 -0.0986 0.9103 0.0438
-4.750 -0.0882 0.01612 0.00816 -0.0998 0.9071 0.0456
-4.500 -0.0621 0.01551 0.00749 -0.0995 0.9009 0.0467
-4.250 -0.0334 0.01479 0.00677 -0.0998 0.8958 0.0483
-4.000 -0.0015 0.01426 0.00621 -0.1007 0.8917 0.0501
-3.750 0.0240 0.01392 0.00584 -0.1002 0.8847 0.0523
-3.500 0.0530 0.01357 0.00544 -0.1004 0.8791 0.0553
-3.250 0.0821 0.01323 0.00503 -0.1006 0.8736 0.0580
-3.000 0.1078 0.01286 0.00467 -0.1002 0.8661 0.0620
-2.750 0.1381 0.01258 0.00432 -0.1005 0.8606 0.0677
-2.500 0.1628 0.01232 0.00406 -0.0998 0.8516 0.0762
-2.250 0.1918 0.01199 0.00376 -0.1000 0.8443 0.0965
-2.000 0.2154 0.01148 0.00357 -0.0993 0.8343 0.1727
-1.750 0.2407 0.01093 0.00345 -0.0990 0.8260 0.3058
-1.500 0.2660 0.01065 0.00340 -0.0984 0.8172 0.3834
-1.250 0.2913 0.01043 0.00339 -0.0978 0.8085 0.4566
-1.000 0.3166 0.01018 0.00337 -0.0972 0.8000 0.5347
-0.750 0.3396 0.00991 0.00339 -0.0959 0.7903 0.6115
-0.500 0.3637 0.00964 0.00335 -0.0948 0.7816 0.6845
-0.250 0.3870 0.00939 0.00335 -0.0934 0.7715 0.7640
0.000 0.4192 0.00919 0.00334 -0.0937 0.7619 0.8692
0.250 0.4742 0.00908 0.00322 -0.0992 0.7529 0.9639
0.500 0.5099 0.00912 0.00318 -0.1009 0.7416 1.0000
0.750 0.5349 0.00919 0.00316 -0.1002 0.7300 1.0000
1.000 0.5601 0.00926 0.00315 -0.0995 0.7180 1.0000
1.250 0.5852 0.00934 0.00313 -0.0988 0.7045 1.0000
1.500 0.6102 0.00944 0.00314 -0.0981 0.6901 1.0000
1.750 0.6352 0.00956 0.00316 -0.0973 0.6757 1.0000
2.000 0.6601 0.00970 0.00319 -0.0966 0.6614 1.0000
2.500 0.7093 0.01002 0.00336 -0.0950 0.6326 1.0000
2.750 0.7338 0.01019 0.00347 -0.0943 0.6185 1.0000
3.000 0.7579 0.01038 0.00360 -0.0934 0.6029 1.0000
3.250 0.7816 0.01057 0.00373 -0.0925 0.5857 1.0000
3.500 0.8050 0.01078 0.00388 -0.0915 0.5674 1.0000
3.750 0.8284 0.01097 0.00405 -0.0905 0.5491 1.0000
4.000 0.8512 0.01118 0.00421 -0.0895 0.5275 1.0000
4.250 0.8726 0.01145 0.00437 -0.0881 0.4989 1.0000
4.500 0.8924 0.01179 0.00456 -0.0865 0.4645 1.0000
4.750 0.9117 0.01219 0.00481 -0.0849 0.4312 1.0000
5.000 0.9318 0.01258 0.00509 -0.0834 0.4019 1.0000
5.250 0.9511 0.01302 0.00540 -0.0818 0.3695 1.0000
5.500 0.9685 0.01358 0.00578 -0.0800 0.3253 1.0000
5.750 0.9831 0.01432 0.00623 -0.0778 0.2716 1.0000
6.000 0.9994 0.01501 0.00672 -0.0759 0.2350 1.0000
6.250 1.0165 0.01566 0.00724 -0.0742 0.2041 1.0000
6.750 1.0431 0.01749 0.00853 -0.0698 0.1068 1.0000
7.000 1.0570 0.01835 0.00925 -0.0676 0.0805 1.0000
7.250 1.0724 0.01909 0.00992 -0.0657 0.0636 1.0000
7.500 1.0874 0.01975 0.01056 -0.0637 0.0554 1.0000
7.750 1.1026 0.02038 0.01123 -0.0616 0.0517 1.0000
8.000 1.1173 0.02102 0.01192 -0.0596 0.0492 1.0000
8.250 1.1307 0.02175 0.01268 -0.0574 0.0471 1.0000
8.500 1.1420 0.02260 0.01358 -0.0550 0.0451 1.0000
8.750 1.1550 0.02336 0.01442 -0.0529 0.0438 1.0000
9.000 1.1671 0.02419 0.01536 -0.0507 0.0425 1.0000
9.250 1.1783 0.02509 0.01635 -0.0486 0.0414 1.0000
9.500 1.1886 0.02608 0.01742 -0.0464 0.0402 1.0000
9.750 1.1981 0.02716 0.01857 -0.0443 0.0390 1.0000
10.000 1.2063 0.02838 0.01985 -0.0422 0.0378 1.0000
10.250 1.2120 0.02986 0.02137 -0.0400 0.0366 1.0000
10.500 1.2194 0.03132 0.02291 -0.0380 0.0357 1.0000
10.750 1.2305 0.03249 0.02421 -0.0365 0.0347 1.0000
11.000 1.2407 0.03377 0.02561 -0.0350 0.0336 1.0000
11.250 1.2503 0.03514 0.02708 -0.0335 0.0325 1.0000
11.500 1.2591 0.03661 0.02864 -0.0321 0.0313 1.0000
11.750 1.2672 0.03814 0.03024 -0.0308 0.0302 1.0000
12.000 1.2743 0.03988 0.03203 -0.0294 0.0291 1.0000
12.250 1.2820 0.04199 0.03418 -0.0281 0.0280 1.0000
12.500 1.2900 0.04357 0.03595 -0.0270 0.0272 1.0000
12.750 1.2970 0.04532 0.03789 -0.0260 0.0262 1.0000
13.000 1.3028 0.04724 0.03997 -0.0250 0.0252 1.0000
13.250 1.3079 0.04927 0.04214 -0.0242 0.0242 1.0000
13.500 1.3121 0.05142 0.04440 -0.0235 0.0234 1.0000
13.750 1.3155 0.05368 0.04674 -0.0230 0.0227 1.0000
14.000 1.3183 0.05609 0.04920 -0.0225 0.0221 1.0000
14.250 1.3195 0.05887 0.05214 -0.0221 0.0215 1.0000
14.500 1.3183 0.06190 0.05543 -0.0220 0.0208 1.0000
14.750 1.3159 0.06518 0.05895 -0.0221 0.0200 1.0000
15.000 1.3129 0.06866 0.06263 -0.0225 0.0194 1.0000
15.250 1.3089 0.07238 0.06653 -0.0231 0.0189 1.0000
15.500 1.3042 0.07630 0.07061 -0.0241 0.0184 1.0000
15.750 1.2992 0.08037 0.07484 -0.0253 0.0180 1.0000
16.000 1.2937 0.08467 0.07928 -0.0267 0.0177 1.0000
16.250 1.2882 0.08909 0.08382 -0.0284 0.0174 1.0000
16.500 1.2825 0.09367 0.08852 -0.0303 0.0172 1.0000
16.750 1.2760 0.09849 0.09344 -0.0325 0.0169 1.0000
17.000 1.2626 0.10489 0.10005 -0.0356 0.0167 1.0000
17.250 1.2444 0.11262 0.10805 -0.0397 0.0165 1.0000
17.500 1.2258 0.12074 0.11639 -0.0443 0.0164 1.0000
17.750 1.2054 0.12965 0.12553 -0.0497 0.0163 1.0000
18.000 1.1808 0.14015 0.13626 -0.0564 0.0163 1.0000
18.250 1.1556 0.15144 0.14773 -0.0637 0.0164 1.0000
18.500 1.1209 0.16632 0.16278 -0.0733 0.0167 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 546 AIRFOIL (goe546-il)