GOE 395 AIRFOIL (goe395-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 395 AIRFOIL (goe395-il) Reynolds number: 500,000 Max Cl/Cd: 149.87 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe395-il-500000.txt Download as CSV file: xf-goe395-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 395 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2429 0.10139 0.09850 -0.0379 0.8098 0.0214
-8.250 -0.2350 0.09849 0.09558 -0.0406 0.8067 0.0215
-8.000 -0.2254 0.09537 0.09244 -0.0432 0.8039 0.0215
-7.750 -0.2155 0.09087 0.08794 -0.0419 0.8020 0.0220
-7.500 -0.2036 0.08836 0.08542 -0.0427 0.7997 0.0224
-7.250 -0.1920 0.08595 0.08301 -0.0442 0.7973 0.0230
-7.000 -0.1786 0.08334 0.08040 -0.0466 0.7951 0.0238
-6.750 -0.1626 0.08041 0.07744 -0.0500 0.7931 0.0250
-6.500 -0.1319 0.07661 0.07359 -0.0624 0.7908 0.0270
-6.250 -0.1069 0.07280 0.06975 -0.0688 0.7891 0.0271
-6.000 -0.0966 0.06830 0.06527 -0.0692 0.7876 0.0278
-5.750 -0.0784 0.06604 0.06300 -0.0703 0.7860 0.0284
-5.500 -0.0560 0.06357 0.06051 -0.0733 0.7842 0.0294
-5.250 -0.0295 0.06066 0.05757 -0.0780 0.7824 0.0313
-5.000 0.0289 0.05572 0.05250 -0.0942 0.7807 0.0342
-4.750 0.0463 0.05129 0.04806 -0.0964 0.7794 0.0353
-4.500 0.0693 0.04948 0.04622 -0.0977 0.7781 0.0362
-4.250 0.0985 0.04724 0.04393 -0.1011 0.7769 0.0379
-4.000 0.1654 0.04012 0.03658 -0.1179 0.7753 0.0434
-3.750 0.1879 0.03853 0.03497 -0.1185 0.7733 0.0443
-3.500 0.2145 0.03702 0.03345 -0.1198 0.7709 0.0455
-3.250 0.2475 0.03495 0.03131 -0.1231 0.7682 0.0479
-3.000 0.3241 0.02202 0.01761 -0.1404 0.7662 0.0444
-2.750 0.3679 0.01410 0.00854 -0.1462 0.7641 0.0418
-2.500 0.3977 0.01309 0.00723 -0.1467 0.7620 0.0424
-2.250 0.4272 0.01210 0.00595 -0.1471 0.7598 0.0431
-2.000 0.4566 0.01119 0.00487 -0.1474 0.7576 0.0445
-1.750 0.4855 0.01080 0.00443 -0.1475 0.7551 0.0460
-1.500 0.5141 0.01054 0.00414 -0.1476 0.7522 0.0477
-1.250 0.5426 0.01033 0.00389 -0.1475 0.7490 0.0502
-1.000 0.5710 0.01019 0.00368 -0.1474 0.7463 0.0529
-0.750 0.5994 0.01003 0.00352 -0.1474 0.7436 0.0587
-0.500 0.6280 0.00990 0.00344 -0.1474 0.7408 0.0676
-0.250 0.6564 0.00994 0.00349 -0.1473 0.7380 0.0788
0.000 0.6846 0.00999 0.00356 -0.1472 0.7352 0.0868
0.250 0.7127 0.00997 0.00352 -0.1471 0.7324 0.0926
0.500 0.7408 0.00998 0.00353 -0.1471 0.7297 0.0986
0.750 0.7689 0.00998 0.00355 -0.1470 0.7268 0.1032
1.000 0.7971 0.00986 0.00347 -0.1470 0.7233 0.1075
1.250 0.8251 0.00978 0.00343 -0.1469 0.7193 0.1132
1.500 0.8530 0.00974 0.00337 -0.1467 0.7151 0.1180
1.750 0.8809 0.00962 0.00329 -0.1466 0.7102 0.1215
2.000 0.9089 0.00950 0.00323 -0.1465 0.7052 0.1270
2.250 0.9368 0.00943 0.00316 -0.1463 0.7004 0.1323
2.500 0.9647 0.00932 0.00314 -0.1462 0.6948 0.1394
2.750 0.9925 0.00921 0.00309 -0.1461 0.6882 0.1504
3.000 1.0203 0.00909 0.00306 -0.1459 0.6817 0.1795
3.250 1.0481 0.00855 0.00330 -0.1461 0.6731 0.5491
3.500 1.0712 0.00772 0.00323 -0.1449 0.6648 1.0000
3.750 1.0987 0.00771 0.00320 -0.1447 0.6544 1.0000
4.000 1.1260 0.00769 0.00318 -0.1444 0.6381 1.0000
4.250 1.1525 0.00769 0.00312 -0.1440 0.5995 1.0000
4.500 1.1750 0.00815 0.00320 -0.1429 0.5361 1.0000
4.750 1.1985 0.00869 0.00351 -0.1422 0.5003 1.0000
5.000 1.2230 0.00913 0.00382 -0.1416 0.4752 1.0000
5.250 1.2477 0.00951 0.00413 -0.1411 0.4526 1.0000
5.500 1.2719 0.00994 0.00445 -0.1405 0.4256 1.0000
5.750 1.2958 0.01038 0.00478 -0.1400 0.3981 1.0000
6.000 1.3194 0.01084 0.00512 -0.1393 0.3710 1.0000
6.250 1.3423 0.01136 0.00552 -0.1386 0.3392 1.0000
6.500 1.3642 0.01196 0.00595 -0.1378 0.3032 1.0000
6.750 1.3846 0.01270 0.00646 -0.1368 0.2585 1.0000
7.000 1.4012 0.01381 0.00719 -0.1353 0.1918 1.0000
7.250 1.4178 0.01490 0.00798 -0.1338 0.1457 1.0000
7.500 1.4367 0.01571 0.00867 -0.1326 0.1223 1.0000
7.750 1.4563 0.01640 0.00928 -0.1314 0.1052 1.0000
8.000 1.4754 0.01711 0.00990 -0.1302 0.0862 1.0000
8.250 1.4915 0.01806 0.01066 -0.1286 0.0583 1.0000
8.500 1.5035 0.01932 0.01172 -0.1264 0.0322 1.0000
8.750 1.5174 0.02034 0.01270 -0.1243 0.0234 1.0000
9.000 1.5303 0.02132 0.01369 -0.1221 0.0202 1.0000
9.250 1.5408 0.02230 0.01476 -0.1195 0.0187 1.0000
9.500 1.5520 0.02327 0.01581 -0.1172 0.0177 1.0000
9.750 1.5616 0.02438 0.01700 -0.1148 0.0168 1.0000
10.000 1.5696 0.02566 0.01836 -0.1124 0.0160 1.0000
10.250 1.5744 0.02725 0.02006 -0.1099 0.0154 1.0000
10.500 1.5747 0.02932 0.02223 -0.1071 0.0148 1.0000
10.750 1.5770 0.03131 0.02434 -0.1047 0.0143 1.0000
11.000 1.5850 0.03283 0.02595 -0.1029 0.0139 1.0000
11.250 1.5916 0.03452 0.02774 -0.1012 0.0134 1.0000
11.500 1.5968 0.03638 0.02970 -0.0994 0.0130 1.0000
11.750 1.6006 0.03841 0.03182 -0.0976 0.0126 1.0000
12.000 1.6037 0.04053 0.03404 -0.0959 0.0123 1.0000
12.250 1.6060 0.04277 0.03638 -0.0942 0.0120 1.0000
12.500 1.6073 0.04514 0.03883 -0.0925 0.0117 1.0000
12.750 1.6072 0.04766 0.04144 -0.0908 0.0115 1.0000
13.000 1.6055 0.05041 0.04428 -0.0890 0.0112 1.0000
13.250 1.6024 0.05338 0.04735 -0.0869 0.0110 1.0000
13.500 1.5987 0.05664 0.05075 -0.0843 0.0108 1.0000
13.750 1.5992 0.05934 0.05359 -0.0830 0.0107 1.0000
14.000 1.5989 0.06211 0.05652 -0.0819 0.0106 1.0000
14.250 1.5975 0.06511 0.05968 -0.0810 0.0104 1.0000
14.500 1.5951 0.06834 0.06307 -0.0802 0.0103 1.0000
14.750 1.5916 0.07180 0.06671 -0.0796 0.0102 1.0000
15.000 1.5868 0.07551 0.07060 -0.0792 0.0100 1.0000
15.250 1.5807 0.07955 0.07481 -0.0791 0.0099 1.0000
15.500 1.5740 0.08383 0.07925 -0.0794 0.0097 1.0000
15.750 1.5658 0.08847 0.08407 -0.0800 0.0096 1.0000
16.000 1.5566 0.09345 0.08921 -0.0809 0.0094 1.0000
16.250 1.5471 0.09867 0.09460 -0.0822 0.0093 1.0000
16.500 1.5352 0.10450 0.10062 -0.0840 0.0093 1.0000
16.750 1.5232 0.11058 0.10688 -0.0863 0.0092 1.0000
17.000 1.5091 0.11725 0.11374 -0.0891 0.0092 1.0000
17.250 1.4946 0.12425 0.12092 -0.0924 0.0092 1.0000
17.500 1.4782 0.13195 0.12882 -0.0964 0.0092 1.0000
17.750 1.4616 0.14005 0.13711 -0.1011 0.0093 1.0000
18.000 1.4444 0.14873 0.14597 -0.1064 0.0093 1.0000
18.250 1.4251 0.15835 0.15578 -0.1126 0.0094 1.0000
18.500 1.4054 0.16876 0.16637 -0.1197 0.0095 1.0000
18.750 1.3835 0.18068 0.17848 -0.1280 0.0097 1.0000
19.000 1.3579 0.19517 0.19314 -0.1382 0.0099 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 395 AIRFOIL (goe395-il)