GOE 394 AIRFOIL (goe394-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 394 AIRFOIL (goe394-il) Reynolds number: 200,000 Max Cl/Cd: 85.88 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe394-il-200000-n5.txt Download as CSV file: xf-goe394-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 394 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2789 0.09714 0.09376 -0.0310 0.9154 0.0287
-7.750 -0.2719 0.09434 0.09093 -0.0336 0.8927 0.0289
-7.250 -0.2546 0.08682 0.08333 -0.0326 0.8668 0.0266
-7.000 -0.2437 0.08331 0.07977 -0.0353 0.8555 0.0253
-6.750 -0.2301 0.07971 0.07611 -0.0389 0.8464 0.0249
-6.500 -0.2131 0.07634 0.07270 -0.0430 0.8386 0.0262
-6.250 -0.1921 0.07236 0.06864 -0.0491 0.8320 0.0277
-6.000 -0.1707 0.06832 0.06454 -0.0543 0.8265 0.0276
-5.750 -0.1450 0.06388 0.06000 -0.0608 0.8213 0.0279
-5.500 -0.1168 0.05946 0.05550 -0.0677 0.8170 0.0304
-5.250 -0.0988 0.05858 0.05459 -0.0672 0.8132 0.0328
-5.000 -0.0684 0.05449 0.05041 -0.0733 0.8089 0.0330
-4.750 -0.0357 0.05034 0.04614 -0.0795 0.8052 0.0333
-4.500 -0.0026 0.04690 0.04256 -0.0848 0.8014 0.0355
-4.250 0.0352 0.04241 0.03787 -0.0915 0.7967 0.0363
-4.000 0.0736 0.03779 0.03301 -0.0977 0.7910 0.0366
-3.750 0.1144 0.03255 0.02740 -0.1040 0.7860 0.0384
-3.500 0.1552 0.02611 0.02045 -0.1099 0.7811 0.0388
-3.250 0.1937 0.01863 0.01186 -0.1147 0.7766 0.0394
-3.000 0.2230 0.01671 0.00944 -0.1154 0.7722 0.0401
-2.750 0.2513 0.01558 0.00802 -0.1156 0.7682 0.0418
-2.500 0.2798 0.01487 0.00714 -0.1158 0.7634 0.0429
-2.250 0.3079 0.01424 0.00631 -0.1157 0.7583 0.0436
-2.000 0.3358 0.01376 0.00565 -0.1155 0.7541 0.0446
-1.750 0.3641 0.01335 0.00515 -0.1155 0.7493 0.0459
-1.500 0.3922 0.01302 0.00473 -0.1154 0.7449 0.0476
-1.250 0.4201 0.01275 0.00436 -0.1152 0.7409 0.0498
-1.000 0.4481 0.01255 0.00417 -0.1150 0.7366 0.0535
-0.750 0.4761 0.01245 0.00412 -0.1149 0.7316 0.0607
-0.500 0.5040 0.01250 0.00417 -0.1147 0.7268 0.0733
-0.250 0.5318 0.01266 0.00424 -0.1144 0.7220 0.0864
0.000 0.5597 0.01278 0.00430 -0.1142 0.7162 0.0958
0.250 0.5872 0.01274 0.00425 -0.1140 0.7107 0.1009
0.500 0.6148 0.01275 0.00424 -0.1138 0.7047 0.1072
0.750 0.6423 0.01268 0.00414 -0.1136 0.6980 0.1101
1.000 0.6696 0.01256 0.00407 -0.1134 0.6917 0.1149
1.250 0.6971 0.01254 0.00408 -0.1132 0.6835 0.1220
1.500 0.7243 0.01239 0.00399 -0.1130 0.6757 0.1267
1.750 0.7517 0.01228 0.00393 -0.1127 0.6663 0.1284
2.000 0.7790 0.01219 0.00390 -0.1125 0.6555 0.1302
2.250 0.8062 0.01211 0.00387 -0.1123 0.6432 0.1322
2.500 0.8334 0.01204 0.00385 -0.1120 0.6277 0.1344
2.750 0.8604 0.01199 0.00380 -0.1116 0.6078 0.1366
3.000 0.8870 0.01196 0.00374 -0.1112 0.5820 0.1399
3.250 0.9125 0.01202 0.00368 -0.1105 0.5516 0.1453
3.500 0.9373 0.01224 0.00374 -0.1099 0.5226 0.1521
3.750 0.9622 0.01249 0.00391 -0.1092 0.4997 0.1642
4.000 0.9874 0.01267 0.00416 -0.1088 0.4826 0.2186
4.250 1.0134 0.01180 0.00446 -0.1086 0.4682 1.0000
4.500 1.0389 0.01214 0.00471 -0.1081 0.4529 1.0000
4.750 1.0643 0.01247 0.00498 -0.1077 0.4392 1.0000
5.250 1.1152 0.01310 0.00558 -0.1068 0.4139 1.0000
5.500 1.1406 0.01341 0.00590 -0.1064 0.4015 1.0000
5.750 1.1655 0.01374 0.00623 -0.1059 0.3869 1.0000
6.000 1.1904 0.01408 0.00660 -0.1054 0.3730 1.0000
6.250 1.2151 0.01442 0.00697 -0.1049 0.3597 1.0000
6.500 1.2397 0.01477 0.00736 -0.1043 0.3461 1.0000
6.750 1.2636 0.01516 0.00778 -0.1037 0.3276 1.0000
7.000 1.2849 0.01576 0.00826 -0.1029 0.2896 1.0000
7.250 1.3017 0.01683 0.00894 -0.1016 0.2195 1.0000
7.500 1.3153 0.01834 0.00996 -0.1000 0.1490 1.0000
7.750 1.3326 0.01941 0.01087 -0.0988 0.1219 1.0000
8.000 1.3508 0.02033 0.01174 -0.0976 0.1048 1.0000
8.250 1.3698 0.02113 0.01257 -0.0965 0.0883 1.0000
8.500 1.3862 0.02215 0.01348 -0.0951 0.0613 1.0000
8.750 1.3979 0.02358 0.01470 -0.0932 0.0366 1.0000
9.000 1.4102 0.02487 0.01599 -0.0913 0.0280 1.0000
9.250 1.4223 0.02609 0.01728 -0.0893 0.0237 1.0000
9.500 1.4329 0.02728 0.01855 -0.0872 0.0208 1.0000
9.750 1.4384 0.02877 0.02012 -0.0845 0.0187 1.0000
10.000 1.4458 0.03014 0.02167 -0.0822 0.0178 1.0000
10.250 1.4521 0.03166 0.02334 -0.0801 0.0169 1.0000
10.500 1.4573 0.03334 0.02516 -0.0781 0.0160 1.0000
10.750 1.4619 0.03516 0.02710 -0.0763 0.0152 1.0000
11.000 1.4655 0.03714 0.02920 -0.0746 0.0144 1.0000
11.250 1.4661 0.03950 0.03166 -0.0731 0.0137 1.0000
11.500 1.4618 0.04247 0.03474 -0.0715 0.0130 1.0000
11.750 1.4633 0.04492 0.03732 -0.0702 0.0127 1.0000
12.000 1.4646 0.04746 0.04001 -0.0690 0.0124 1.0000
12.250 1.4651 0.05013 0.04283 -0.0679 0.0120 1.0000
12.500 1.4649 0.05292 0.04577 -0.0669 0.0117 1.0000
12.750 1.4642 0.05583 0.04886 -0.0659 0.0115 1.0000
13.000 1.4630 0.05884 0.05202 -0.0651 0.0112 1.0000
13.250 1.4613 0.06198 0.05531 -0.0644 0.0109 1.0000
13.500 1.4591 0.06525 0.05872 -0.0638 0.0107 1.0000
13.750 1.4561 0.06866 0.06228 -0.0635 0.0104 1.0000
14.000 1.4524 0.07227 0.06603 -0.0635 0.0101 1.0000
14.250 1.4484 0.07606 0.06996 -0.0638 0.0099 1.0000
14.500 1.4433 0.08011 0.07413 -0.0643 0.0097 1.0000
14.750 1.4375 0.08440 0.07855 -0.0650 0.0095 1.0000
15.000 1.4302 0.08901 0.08329 -0.0658 0.0092 1.0000
15.250 1.4211 0.09405 0.08846 -0.0668 0.0091 1.0000
15.500 1.4129 0.09931 0.09394 -0.0686 0.0090 1.0000
15.750 1.4037 0.10490 0.09974 -0.0707 0.0089 1.0000
16.000 1.3937 0.11081 0.10585 -0.0731 0.0088 1.0000
16.250 1.3826 0.11713 0.11238 -0.0760 0.0088 1.0000
16.500 1.3707 0.12375 0.11920 -0.0792 0.0087 1.0000
16.750 1.3582 0.13078 0.12642 -0.0829 0.0087 1.0000
17.000 1.3452 0.13818 0.13401 -0.0871 0.0087 1.0000
17.250 1.3319 0.14601 0.14202 -0.0917 0.0087 1.0000
17.500 1.3182 0.15434 0.15053 -0.0969 0.0087 1.0000
17.750 1.3039 0.16327 0.15962 -0.1026 0.0087 1.0000
18.000 1.2893 0.17287 0.16938 -0.1089 0.0088 1.0000
18.250 1.2740 0.18343 0.18007 -0.1158 0.0089 1.0000
18.500 1.2571 0.19553 0.19230 -0.1237 0.0090 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 394 AIRFOIL (goe394-il)