GOE 507 AIRFOIL (goe507-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 507 AIRFOIL (goe507-il) Reynolds number: 50,000 Max Cl/Cd: 37.98 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe507-il-50000-n5.txt Download as CSV file: xf-goe507-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 507 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4298 0.11047 0.10296 -0.0294 1.0000 0.0763
-9.750 -0.4290 0.10673 0.09926 -0.0307 1.0000 0.0762
-9.500 -0.4291 0.10300 0.09559 -0.0321 1.0000 0.0762
-9.250 -0.4307 0.09922 0.09188 -0.0336 1.0000 0.0763
-9.000 -0.4312 0.09554 0.08826 -0.0347 1.0000 0.0759
-8.750 -0.4333 0.09185 0.08464 -0.0357 1.0000 0.0753
-8.500 -0.4394 0.08821 0.08109 -0.0369 1.0000 0.0754
-8.250 -0.4484 0.08467 0.07767 -0.0380 1.0000 0.0756
-8.000 -0.4555 0.08069 0.07377 -0.0397 1.0000 0.0759
-7.750 -0.4631 0.07645 0.06958 -0.0415 1.0000 0.0763
-7.500 -0.4706 0.07191 0.06506 -0.0431 1.0000 0.0768
-7.250 -0.4770 0.06728 0.06038 -0.0441 1.0000 0.0771
-7.000 -0.4825 0.06246 0.05543 -0.0447 1.0000 0.0774
-6.500 -0.4759 0.05723 0.05009 -0.0423 1.0000 0.0823
-6.250 -0.4718 0.05392 0.04662 -0.0416 1.0000 0.0845
-6.000 -0.4667 0.04988 0.04227 -0.0414 1.0000 0.0862
-5.750 -0.4595 0.04556 0.03740 -0.0414 1.0000 0.0897
-5.500 -0.4489 0.04170 0.03299 -0.0410 1.0000 0.0925
-5.250 -0.4343 0.03988 0.03105 -0.0398 1.0000 0.0947
-5.000 -0.4186 0.03838 0.02938 -0.0387 1.0000 0.0989
-4.750 -0.4015 0.03602 0.02653 -0.0380 1.0000 0.1027
-4.500 -0.3825 0.03362 0.02356 -0.0373 1.0000 0.1046
-4.250 -0.3624 0.03169 0.02115 -0.0365 1.0000 0.1066
-4.000 -0.3435 0.03082 0.02027 -0.0356 1.0000 0.1104
-3.750 -0.3232 0.02985 0.01909 -0.0347 1.0000 0.1146
-3.500 -0.3012 0.02863 0.01751 -0.0340 1.0000 0.1170
-3.250 -0.2787 0.02753 0.01609 -0.0333 1.0000 0.1191
-3.000 -0.2498 0.02658 0.01483 -0.0337 0.9974 0.1216
-2.750 -0.2146 0.02587 0.01405 -0.0355 0.9914 0.1251
-2.500 -0.1771 0.02532 0.01335 -0.0376 0.9853 0.1299
-2.250 -0.1403 0.02483 0.01261 -0.0394 0.9783 0.1366
-2.000 -0.1042 0.02455 0.01234 -0.0414 0.9711 0.1478
-1.750 -0.0669 0.02426 0.01202 -0.0434 0.9642 0.1624
-1.500 -0.0314 0.02388 0.01160 -0.0451 0.9570 0.1826
-1.250 0.0024 0.02344 0.01127 -0.0465 0.9495 0.2170
-1.000 0.0396 0.02299 0.01104 -0.0487 0.9430 0.2723
-0.750 0.0693 0.02253 0.01095 -0.0495 0.9342 0.3444
-0.500 0.1295 0.02069 0.01102 -0.0552 0.9317 1.0000
-0.250 0.1618 0.02092 0.01098 -0.0562 0.9221 1.0000
0.000 0.1998 0.02114 0.01097 -0.0583 0.9139 1.0000
0.250 0.2325 0.02135 0.01102 -0.0594 0.9037 1.0000
0.500 0.2646 0.02156 0.01110 -0.0603 0.8933 1.0000
0.750 0.3063 0.02171 0.01113 -0.0629 0.8853 1.0000
1.000 0.3352 0.02190 0.01125 -0.0631 0.8733 1.0000
1.250 0.3662 0.02206 0.01137 -0.0636 0.8618 1.0000
1.500 0.4010 0.02216 0.01143 -0.0647 0.8513 1.0000
1.750 0.4349 0.02222 0.01150 -0.0655 0.8401 1.0000
2.000 0.4630 0.02234 0.01162 -0.0652 0.8266 1.0000
2.250 0.4916 0.02243 0.01173 -0.0650 0.8127 1.0000
2.500 0.5209 0.02247 0.01181 -0.0648 0.7986 1.0000
2.750 0.5502 0.02246 0.01186 -0.0644 0.7839 1.0000
3.000 0.5790 0.02242 0.01188 -0.0639 0.7682 1.0000
3.250 0.6073 0.02236 0.01189 -0.0632 0.7518 1.0000
3.500 0.6352 0.02229 0.01189 -0.0624 0.7348 1.0000
3.750 0.6629 0.02222 0.01192 -0.0616 0.7175 1.0000
4.000 0.6906 0.02216 0.01195 -0.0607 0.6999 1.0000
4.250 0.7144 0.02227 0.01217 -0.0594 0.6800 1.0000
4.500 0.7392 0.02235 0.01236 -0.0582 0.6598 1.0000
4.750 0.7659 0.02235 0.01246 -0.0571 0.6393 1.0000
5.000 0.7899 0.02240 0.01260 -0.0556 0.6139 1.0000
5.250 0.8145 0.02238 0.01260 -0.0540 0.5848 1.0000
5.500 0.8374 0.02242 0.01258 -0.0520 0.5494 1.0000
5.750 0.8591 0.02263 0.01268 -0.0500 0.5125 1.0000
6.000 0.8766 0.02308 0.01307 -0.0477 0.4723 1.0000
6.250 0.8927 0.02361 0.01351 -0.0453 0.4305 1.0000
6.500 0.9085 0.02422 0.01409 -0.0431 0.3910 1.0000
6.750 0.9239 0.02492 0.01479 -0.0409 0.3516 1.0000
7.000 0.9367 0.02579 0.01556 -0.0385 0.3057 1.0000
7.250 0.9459 0.02700 0.01651 -0.0357 0.2504 1.0000
7.500 0.9532 0.02855 0.01773 -0.0331 0.1955 1.0000
7.750 0.9618 0.03019 0.01910 -0.0307 0.1590 1.0000
8.000 0.9722 0.03179 0.02057 -0.0286 0.1376 1.0000
8.250 0.9828 0.03337 0.02207 -0.0266 0.1213 1.0000
8.750 1.0055 0.03662 0.02532 -0.0227 0.0986 1.0000
9.000 1.0203 0.03841 0.02719 -0.0212 0.0892 1.0000
9.250 1.0368 0.04038 0.02917 -0.0199 0.0811 1.0000
9.500 1.0554 0.04235 0.03129 -0.0189 0.0735 1.0000
9.750 1.0779 0.04473 0.03384 -0.0183 0.0672 1.0000
10.000 1.0958 0.04702 0.03629 -0.0174 0.0619 1.0000
10.250 1.1150 0.04988 0.03935 -0.0168 0.0581 1.0000
10.500 1.1217 0.05259 0.04255 -0.0147 0.0550 1.0000
10.750 1.1264 0.05532 0.04559 -0.0127 0.0524 1.0000
11.000 1.1295 0.05821 0.04872 -0.0109 0.0506 1.0000
11.250 1.1319 0.06115 0.05182 -0.0093 0.0488 1.0000
11.500 1.1336 0.06466 0.05540 -0.0081 0.0474 1.0000
11.750 1.1197 0.06810 0.05921 -0.0061 0.0469 1.0000
12.000 1.1020 0.07201 0.06347 -0.0048 0.0464 1.0000
12.250 1.0841 0.07637 0.06811 -0.0046 0.0462 1.0000
12.500 1.0648 0.08132 0.07332 -0.0053 0.0462 1.0000
12.750 1.0445 0.08689 0.07910 -0.0070 0.0463 1.0000
13.000 1.0212 0.09339 0.08579 -0.0099 0.0464 1.0000
13.250 0.9996 0.10049 0.09304 -0.0137 0.0466 1.0000
13.500 0.9777 0.10848 0.10114 -0.0184 0.0470 1.0000
13.750 0.9582 0.11690 0.10964 -0.0235 0.0473 1.0000
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Polar data table (+)
Polar graphs
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