GOE 422 AIRFOIL (goe422-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 422 AIRFOIL (goe422-il) Reynolds number: 100,000 Max Cl/Cd: 44.91 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe422-il-100000.txt Download as CSV file: xf-goe422-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 422 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.0528 0.10724 0.10235 -0.0873 0.9134 0.1215
-8.000 -0.0690 0.10586 0.10101 -0.0897 0.9046 0.1254
-7.750 -0.1238 0.10572 0.10094 -0.0939 0.8912 0.1266
-7.500 -0.0414 0.09874 0.09391 -0.0934 0.8947 0.1300
-7.250 -0.0442 0.09756 0.09276 -0.0904 0.8851 0.1324
-7.000 -0.0391 0.09497 0.09018 -0.0927 0.8793 0.1374
-6.750 -0.0807 0.09471 0.08999 -0.0912 0.8652 0.1400
-6.500 -0.0667 0.08969 0.08495 -0.0963 0.8610 0.1432
-6.250 -0.0763 0.08904 0.08436 -0.0908 0.8501 0.1444
-6.000 -0.0516 0.08617 0.08146 -0.0920 0.8461 0.1487
-5.750 -0.0639 0.08460 0.07991 -0.0917 0.8357 0.1532
-5.500 -0.0649 0.08029 0.07551 -0.0973 0.8278 0.1595
-5.250 -0.0298 0.07741 0.07264 -0.0980 0.8256 0.1641
-5.000 -0.0501 0.07713 0.07239 -0.0933 0.8125 0.1666
-4.750 -0.0281 0.07266 0.06777 -0.0997 0.8080 0.1779
-4.500 -0.0340 0.07209 0.06724 -0.0954 0.7974 0.1804
-4.250 -0.0145 0.06880 0.06378 -0.0999 0.7910 0.1950
-4.000 0.0208 0.06576 0.06059 -0.1042 0.7880 0.2130
-3.750 0.0601 0.06303 0.05793 -0.1050 0.7865 0.2235
-3.500 0.0689 0.04908 0.04197 -0.1147 0.7726 0.1151
-3.250 0.1129 0.04574 0.03822 -0.1176 0.7697 0.1097
-3.000 0.1618 0.04257 0.03454 -0.1210 0.7679 0.1083
-2.750 0.1600 0.04292 0.03472 -0.1172 0.7545 0.1091
-2.500 0.2051 0.04090 0.03221 -0.1193 0.7512 0.1120
-2.250 0.2563 0.03845 0.02939 -0.1222 0.7494 0.1148
-2.000 0.2595 0.03888 0.02981 -0.1190 0.7365 0.1167
-1.750 0.3036 0.03730 0.02809 -0.1207 0.7330 0.1239
-1.500 0.3547 0.03519 0.02592 -0.1234 0.7313 0.1341
-1.250 0.3720 0.03520 0.02582 -0.1217 0.7222 0.1420
-1.000 0.4537 0.03179 0.02254 -0.1285 0.7270 0.1732
-0.750 0.4509 0.03259 0.02349 -0.1241 0.7135 0.1836
-0.500 0.5014 0.03101 0.02207 -0.1269 0.7116 0.2236
-0.250 0.5364 0.03015 0.02145 -0.1275 0.7070 0.2745
0.000 0.5445 0.03032 0.02206 -0.1247 0.6958 0.3545
0.250 0.5936 0.02756 0.02098 -0.1254 0.6939 1.0000
0.500 0.6448 0.02661 0.01967 -0.1283 0.6915 1.0000
0.750 0.6415 0.02786 0.02085 -0.1238 0.6783 1.0000
1.000 0.6893 0.02696 0.01969 -0.1261 0.6750 1.0000
1.250 0.7412 0.02598 0.01845 -0.1292 0.6724 1.0000
1.500 0.7370 0.02724 0.01969 -0.1245 0.6592 1.0000
1.750 0.7857 0.02634 0.01858 -0.1271 0.6558 1.0000
2.000 0.7945 0.02711 0.01930 -0.1242 0.6451 1.0000
2.250 0.8351 0.02650 0.01852 -0.1257 0.6395 1.0000
2.500 0.8876 0.02551 0.01732 -0.1290 0.6359 1.0000
2.750 0.8886 0.02651 0.01832 -0.1248 0.6236 1.0000
3.000 0.9367 0.02572 0.01735 -0.1275 0.6189 1.0000
3.250 0.9459 0.02648 0.01810 -0.1246 0.6084 1.0000
3.500 0.9861 0.02603 0.01751 -0.1262 0.6021 1.0000
3.750 1.0104 0.02627 0.01767 -0.1256 0.5939 1.0000
4.000 1.0368 0.02635 0.01767 -0.1252 0.5853 1.0000
4.250 1.0825 0.02587 0.01700 -0.1278 0.5796 1.0000
4.500 1.0873 0.02676 0.01794 -0.1242 0.5686 1.0000
4.750 1.1349 0.02629 0.01727 -0.1271 0.5631 1.0000
5.000 1.1359 0.02744 0.01850 -0.1231 0.5528 1.0000
5.250 1.1759 0.02722 0.01813 -0.1249 0.5465 1.0000
5.500 1.1876 0.02806 0.01900 -0.1226 0.5379 1.0000
5.750 1.2140 0.02830 0.01918 -0.1223 0.5304 1.0000
6.000 1.2496 0.02836 0.01910 -0.1236 0.5239 1.0000
6.250 1.2531 0.02936 0.02019 -0.1200 0.5151 1.0000
6.500 1.2992 0.02898 0.01961 -0.1227 0.5089 1.0000
6.750 1.2978 0.03025 0.02101 -0.1185 0.5010 1.0000
7.000 1.3210 0.03063 0.02135 -0.1179 0.4944 1.0000
7.250 1.3674 0.03045 0.02099 -0.1208 0.4890 1.0000
7.500 1.3522 0.03217 0.02293 -0.1147 0.4820 1.0000
7.750 1.3741 0.03266 0.02341 -0.1140 0.4764 1.0000
8.000 1.4272 0.03235 0.02291 -0.1179 0.4718 1.0000
8.250 1.4024 0.03437 0.02519 -0.1105 0.4657 1.0000
8.500 1.4092 0.03546 0.02634 -0.1079 0.4602 1.0000
8.750 1.4476 0.03550 0.02632 -0.1096 0.4561 1.0000
9.000 1.4742 0.03610 0.02691 -0.1098 0.4518 1.0000
9.250 1.4324 0.03930 0.03038 -0.1011 0.4465 1.0000
9.500 1.4401 0.04053 0.03168 -0.0990 0.4415 1.0000
9.750 1.5005 0.03925 0.03026 -0.1029 0.4373 1.0000
10.000 1.4900 0.04139 0.03254 -0.0986 0.4328 1.0000
10.250 1.2443 0.06402 0.05576 -0.0802 0.4196 1.0000
10.500 1.4026 0.05083 0.04239 -0.0859 0.4218 1.0000
10.750 1.5126 0.04531 0.03668 -0.0927 0.4194 1.0000
11.750 1.5754 0.04838 0.03992 -0.0883 0.4002 1.0000
12.000 0.9348 0.13192 0.12427 -0.0828 0.3590 1.0000
12.250 0.9516 0.13286 0.12523 -0.0824 0.3529 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 422 AIRFOIL (goe422-il)