GOE 362 AIRFOIL (goe362-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 362 AIRFOIL (goe362-il) Reynolds number: 200,000 Max Cl/Cd: 79.57 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe362-il-200000.txt Download as CSV file: xf-goe362-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 362 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3290 0.10395 0.10042 -0.0250 1.0000 0.0316
-8.250 -0.3294 0.10211 0.09865 -0.0264 1.0000 0.0318
-8.000 -0.3300 0.10011 0.09671 -0.0268 1.0000 0.0319
-7.750 -0.3290 0.09783 0.09449 -0.0281 1.0000 0.0319
-7.500 -0.3273 0.09311 0.08984 -0.0278 1.0000 0.0323
-7.250 -0.3200 0.08885 0.08558 -0.0242 1.0000 0.0331
-7.000 -0.3149 0.08609 0.08286 -0.0230 1.0000 0.0338
-6.750 -0.3124 0.08375 0.08057 -0.0224 1.0000 0.0345
-6.500 -0.3113 0.08156 0.07843 -0.0222 1.0000 0.0351
-6.250 -0.3110 0.07940 0.07633 -0.0218 1.0000 0.0360
-6.000 -0.3113 0.07730 0.07428 -0.0216 1.0000 0.0369
-5.750 -0.3116 0.07522 0.07225 -0.0216 1.0000 0.0376
-5.500 -0.3109 0.07305 0.07012 -0.0220 1.0000 0.0387
-5.250 -0.2678 0.06883 0.06580 -0.0366 0.9968 0.0418
-5.000 -0.2187 0.06134 0.05819 -0.0498 0.9909 0.0429
-4.750 -0.1949 0.05754 0.05442 -0.0509 0.9857 0.0446
-4.500 -0.1562 0.05381 0.05062 -0.0567 0.9796 0.0486
-4.250 -0.0912 0.04678 0.04328 -0.0720 0.9734 0.0557
-4.000 -0.0574 0.04380 0.04030 -0.0751 0.9688 0.0587
-3.750 0.0014 0.03833 0.03431 -0.0856 0.9615 0.0680
-3.500 0.0359 0.03525 0.03130 -0.0887 0.9574 0.0708
-3.250 0.0839 0.03137 0.02694 -0.0945 0.9506 0.0818
-3.000 0.1314 0.02385 0.01859 -0.0984 0.9452 0.0535
-2.750 0.1721 0.01907 0.01320 -0.1012 0.9404 0.0479
-2.500 0.2066 0.01653 0.01010 -0.1019 0.9306 0.0482
-2.250 0.2412 0.01543 0.00865 -0.1026 0.9210 0.0535
-2.000 0.2755 0.01380 0.00681 -0.1033 0.9122 0.0581
-1.750 0.3053 0.01331 0.00617 -0.1030 0.8990 0.0667
-1.500 0.3338 0.01253 0.00539 -0.1026 0.8853 0.0765
-1.250 0.3619 0.01209 0.00489 -0.1020 0.8709 0.0918
-1.000 0.3890 0.01150 0.00432 -0.1013 0.8560 0.1142
-0.750 0.4156 0.01102 0.00389 -0.1006 0.8406 0.1470
-0.500 0.4422 0.01084 0.00370 -0.0999 0.8248 0.1804
-0.250 0.4681 0.01062 0.00355 -0.0992 0.8079 0.2156
0.000 0.4932 0.01021 0.00342 -0.0984 0.7908 0.2988
0.250 0.5262 0.00862 0.00326 -0.0991 0.7741 1.0000
0.500 0.5525 0.00873 0.00315 -0.0983 0.7564 1.0000
0.750 0.5786 0.00886 0.00308 -0.0976 0.7386 1.0000
1.000 0.6046 0.00900 0.00304 -0.0968 0.7200 1.0000
1.250 0.6305 0.00914 0.00303 -0.0961 0.7003 1.0000
1.500 0.6564 0.00932 0.00304 -0.0953 0.6810 1.0000
1.750 0.6821 0.00951 0.00308 -0.0946 0.6609 1.0000
2.000 0.7078 0.00971 0.00315 -0.0939 0.6404 1.0000
2.250 0.7334 0.00994 0.00323 -0.0932 0.6210 1.0000
2.500 0.7591 0.01016 0.00336 -0.0926 0.6012 1.0000
2.750 0.7849 0.01039 0.00349 -0.0920 0.5822 1.0000
3.000 0.8106 0.01063 0.00362 -0.0914 0.5646 1.0000
3.250 0.8363 0.01088 0.00378 -0.0909 0.5482 1.0000
3.500 0.8620 0.01113 0.00397 -0.0903 0.5327 1.0000
3.750 0.8878 0.01139 0.00416 -0.0898 0.5180 1.0000
4.000 0.9135 0.01165 0.00438 -0.0893 0.5038 1.0000
4.250 0.9392 0.01193 0.00462 -0.0888 0.4904 1.0000
4.500 0.9647 0.01222 0.00491 -0.0883 0.4769 1.0000
4.750 0.9898 0.01252 0.00517 -0.0877 0.4621 1.0000
5.000 1.0145 0.01284 0.00543 -0.0870 0.4462 1.0000
5.250 1.0392 0.01317 0.00572 -0.0864 0.4307 1.0000
5.500 1.0637 0.01347 0.00605 -0.0857 0.4150 1.0000
5.750 1.0882 0.01378 0.00637 -0.0851 0.4002 1.0000
6.000 1.1126 0.01408 0.00671 -0.0844 0.3856 1.0000
6.250 1.1369 0.01436 0.00705 -0.0838 0.3709 1.0000
6.500 1.1610 0.01463 0.00742 -0.0831 0.3552 1.0000
6.750 1.1848 0.01489 0.00776 -0.0823 0.3368 1.0000
7.000 1.2078 0.01520 0.00810 -0.0815 0.3161 1.0000
7.250 1.2305 0.01553 0.00848 -0.0806 0.2912 1.0000
7.500 1.2514 0.01604 0.00896 -0.0795 0.2579 1.0000
7.750 1.2687 0.01697 0.00965 -0.0780 0.2009 1.0000
8.000 1.2817 0.01850 0.01075 -0.0762 0.1389 1.0000
8.250 1.2965 0.01982 0.01187 -0.0745 0.1112 1.0000
8.500 1.3116 0.02105 0.01302 -0.0729 0.0962 1.0000
8.750 1.3275 0.02214 0.01417 -0.0713 0.0853 1.0000
9.000 1.3425 0.02328 0.01538 -0.0696 0.0754 1.0000
9.250 1.3548 0.02462 0.01676 -0.0676 0.0661 1.0000
9.500 1.3649 0.02609 0.01824 -0.0654 0.0558 1.0000
9.750 1.3721 0.02773 0.01988 -0.0629 0.0462 1.0000
10.000 1.3824 0.02895 0.02125 -0.0605 0.0398 1.0000
10.250 1.3820 0.03108 0.02335 -0.0571 0.0361 1.0000
10.500 1.3897 0.03253 0.02501 -0.0546 0.0332 1.0000
10.750 1.3960 0.03398 0.02656 -0.0523 0.0305 1.0000
11.000 1.3992 0.03581 0.02844 -0.0502 0.0284 1.0000
11.250 1.4029 0.03866 0.03138 -0.0480 0.0269 1.0000
11.500 1.4098 0.04072 0.03365 -0.0462 0.0258 1.0000
11.750 1.4158 0.04306 0.03620 -0.0445 0.0248 1.0000
12.000 1.4203 0.04556 0.03890 -0.0429 0.0238 1.0000
12.250 1.4226 0.04819 0.04172 -0.0416 0.0230 1.0000
12.500 1.4229 0.05074 0.04441 -0.0405 0.0222 1.0000
12.750 1.4227 0.05348 0.04723 -0.0398 0.0212 1.0000
13.000 1.4199 0.05752 0.05144 -0.0390 0.0206 1.0000
13.250 1.4090 0.06280 0.05703 -0.0386 0.0202 1.0000
13.500 1.3988 0.06703 0.06152 -0.0389 0.0201 1.0000
13.750 1.3863 0.07181 0.06656 -0.0397 0.0200 1.0000
14.000 1.3725 0.07714 0.07214 -0.0411 0.0200 1.0000
14.250 1.3567 0.08297 0.07820 -0.0432 0.0199 1.0000
14.500 1.3391 0.08938 0.08485 -0.0460 0.0200 1.0000
14.750 1.3210 0.09627 0.09195 -0.0493 0.0200 1.0000
15.000 1.3019 0.10374 0.09962 -0.0534 0.0201 1.0000
15.250 1.2821 0.11183 0.10790 -0.0580 0.0202 1.0000
15.500 1.2649 0.11982 0.11606 -0.0630 0.0203 1.0000
15.750 1.2494 0.12788 0.12428 -0.0685 0.0204 1.0000
16.000 1.2320 0.13710 0.13368 -0.0749 0.0206 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 362 AIRFOIL (goe362-il)