GOE 265 AIRFOIL (goe265-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 265 AIRFOIL (goe265-il) Reynolds number: 50,000 Max Cl/Cd: 39.07 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe265-il-50000-n5.txt Download as CSV file: xf-goe265-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 265 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.2176 0.10990 0.10390 -0.0259 1.0000 0.0760
-7.500 -0.2261 0.11011 0.10431 -0.0260 1.0000 0.0764
-7.250 -0.2377 0.11044 0.10481 -0.0252 1.0000 0.0766
-7.000 -0.2253 0.10549 0.09992 -0.0260 0.9939 0.0774
-6.750 -0.2022 0.09935 0.09376 -0.0269 0.9853 0.0795
-6.500 -0.1758 0.09516 0.08955 -0.0319 0.9717 0.0822
-6.250 -0.1469 0.09149 0.08583 -0.0387 0.9572 0.0861
-6.000 -0.1019 0.09044 0.08466 -0.0555 0.9384 0.0903
-5.750 -0.0816 0.08458 0.07883 -0.0563 0.9286 0.0920
-5.500 -0.0576 0.08011 0.07432 -0.0581 0.9180 0.0950
-5.250 -0.0301 0.07676 0.07090 -0.0628 0.9043 0.0989
-5.000 0.0218 0.07568 0.06956 -0.0781 0.8884 0.1053
-4.750 0.0333 0.07069 0.06463 -0.0759 0.8768 0.1075
-4.500 0.0552 0.06738 0.06127 -0.0772 0.8646 0.1117
-4.250 0.0887 0.06472 0.05845 -0.0828 0.8524 0.1177
-4.000 0.1258 0.06204 0.05560 -0.0897 0.8397 0.1219
-3.750 0.1419 0.05910 0.05264 -0.0888 0.8272 0.1279
-3.250 0.2107 0.05403 0.04719 -0.0980 0.8042 0.1378
-2.500 0.3229 0.04545 0.03774 -0.1100 0.7713 0.1038
-2.250 0.3526 0.04298 0.03508 -0.1115 0.7607 0.0999
-2.000 0.3867 0.04107 0.03289 -0.1139 0.7489 0.1018
-1.750 0.4204 0.03911 0.03063 -0.1159 0.7384 0.1016
-1.500 0.4539 0.03719 0.02840 -0.1174 0.7280 0.1003
-1.250 0.4863 0.03557 0.02647 -0.1188 0.7164 0.1001
-1.000 0.5216 0.03409 0.02452 -0.1203 0.7061 0.1044
-0.750 0.5519 0.03281 0.02298 -0.1209 0.6959 0.1074
-0.500 0.5810 0.03177 0.02171 -0.1213 0.6847 0.1092
-0.250 0.6110 0.03085 0.02051 -0.1216 0.6751 0.1138
0.000 0.6422 0.03002 0.01924 -0.1219 0.6646 0.1211
0.250 0.6692 0.02939 0.01849 -0.1218 0.6539 0.1252
0.500 0.6992 0.02878 0.01749 -0.1216 0.6451 0.1345
0.750 0.7257 0.02842 0.01705 -0.1214 0.6338 0.1419
1.000 0.7548 0.02808 0.01631 -0.1212 0.6236 0.1502
1.250 0.7828 0.02775 0.01582 -0.1208 0.6145 0.1610
1.500 0.8095 0.02758 0.01547 -0.1203 0.6034 0.1678
2.000 0.8629 0.02728 0.01495 -0.1192 0.5842 0.1901
2.250 0.8882 0.02729 0.01489 -0.1185 0.5738 0.1983
2.500 0.9150 0.02715 0.01466 -0.1178 0.5658 0.2132
2.750 0.9392 0.02735 0.01489 -0.1173 0.5549 0.2270
3.000 0.9647 0.02744 0.01496 -0.1166 0.5459 0.2412
3.250 0.9902 0.02751 0.01510 -0.1161 0.5368 0.2660
3.500 1.0149 0.02769 0.01547 -0.1157 0.5274 0.3059
3.750 1.0391 0.02693 0.01551 -0.1147 0.5196 1.0000
4.000 1.0627 0.02759 0.01603 -0.1140 0.5100 1.0000
4.250 1.0890 0.02795 0.01619 -0.1134 0.5029 1.0000
4.500 1.1112 0.02874 0.01697 -0.1127 0.4935 1.0000
4.750 1.1373 0.02911 0.01720 -0.1121 0.4870 1.0000
5.000 1.1585 0.02999 0.01816 -0.1114 0.4778 1.0000
5.250 1.1841 0.03042 0.01851 -0.1108 0.4716 1.0000
5.500 1.2046 0.03140 0.01958 -0.1101 0.4634 1.0000
5.750 1.2289 0.03198 0.02016 -0.1095 0.4571 1.0000
6.000 1.2500 0.03289 0.02114 -0.1088 0.4500 1.0000
6.250 1.2719 0.03369 0.02201 -0.1080 0.4431 1.0000
6.500 1.2956 0.03429 0.02262 -0.1073 0.4372 1.0000
6.750 1.3134 0.03538 0.02388 -0.1063 0.4289 1.0000
7.000 1.3383 0.03567 0.02415 -0.1056 0.4223 1.0000
7.250 1.3540 0.03679 0.02545 -0.1043 0.4131 1.0000
7.500 1.3763 0.03721 0.02592 -0.1033 0.4053 1.0000
7.750 1.3937 0.03802 0.02686 -0.1021 0.3963 1.0000
8.000 1.4114 0.03878 0.02772 -0.1008 0.3877 1.0000
8.250 1.4321 0.03918 0.02817 -0.0997 0.3792 1.0000
8.500 1.4446 0.04039 0.02960 -0.0981 0.3705 1.0000
8.750 1.4664 0.04081 0.03007 -0.0971 0.3634 1.0000
9.000 1.4754 0.04241 0.03190 -0.0954 0.3557 1.0000
9.250 1.4942 0.04309 0.03268 -0.0942 0.3487 1.0000
9.500 1.5015 0.04470 0.03451 -0.0924 0.3409 1.0000
9.750 1.5179 0.04544 0.03539 -0.0909 0.3332 1.0000
10.000 1.5193 0.04739 0.03755 -0.0888 0.3251 1.0000
10.250 1.5356 0.04802 0.03828 -0.0872 0.3172 1.0000
10.500 1.5238 0.05080 0.04128 -0.0843 0.3094 1.0000
10.750 1.5384 0.05151 0.04208 -0.0827 0.3015 1.0000
11.000 1.5152 0.05574 0.04655 -0.0805 0.2938 1.0000
11.250 1.5280 0.05671 0.04761 -0.0791 0.2861 1.0000
11.500 1.4890 0.06353 0.05465 -0.0786 0.2779 1.0000
11.750 1.5056 0.06407 0.05532 -0.0772 0.2705 1.0000
12.000 1.4357 0.07649 0.06789 -0.0798 0.2602 1.0000
12.250 1.4736 0.07380 0.06531 -0.0771 0.2546 1.0000
12.500 1.3854 0.09084 0.08240 -0.0826 0.2420 1.0000
12.750 1.4329 0.08597 0.07770 -0.0789 0.2387 1.0000
13.000 1.3461 0.10459 0.09627 -0.0860 0.2254 1.0000
13.500 1.3034 0.12043 0.11219 -0.0910 0.2103 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 265 AIRFOIL (goe265-il)