GOE 81 AIRFOIL (goe81-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 81 AIRFOIL (goe81-il) Reynolds number: 500,000 Max Cl/Cd: 102.53 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe81-il-500000-n5.txt Download as CSV file: xf-goe81-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 81 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.1692 0.08833 0.08580 -0.0509 0.8862 0.0147
-7.750 -0.1605 0.08551 0.08293 -0.0525 0.8752 0.0147
-7.500 -0.1515 0.08273 0.08012 -0.0543 0.8643 0.0149
-7.250 -0.1384 0.07889 0.07624 -0.0594 0.8536 0.0163
-7.000 -0.1248 0.07676 0.07407 -0.0606 0.8431 0.0167
-6.750 -0.1087 0.07416 0.07144 -0.0634 0.8315 0.0173
-6.500 -0.0888 0.07102 0.06824 -0.0680 0.8201 0.0189
-6.250 -0.0662 0.06737 0.06451 -0.0737 0.8081 0.0191
-6.000 -0.0070 0.04612 0.04321 -0.0784 0.7585 0.0193
-5.750 0.0055 0.04346 0.04049 -0.0782 0.7466 0.0195
-5.500 0.0224 0.04080 0.03777 -0.0800 0.7353 0.0199
-5.250 0.0425 0.03798 0.03489 -0.0829 0.7251 0.0205
-5.000 0.0651 0.03500 0.03182 -0.0864 0.7159 0.0215
-4.750 0.1044 0.03074 0.02736 -0.0954 0.7065 0.0228
-4.500 0.1335 0.02726 0.02373 -0.0997 0.6979 0.0229
-4.250 0.1621 0.02412 0.02043 -0.1031 0.6899 0.0229
-4.000 0.1859 0.02107 0.01729 -0.1053 0.6831 0.0231
-3.750 0.2063 0.01925 0.01541 -0.1060 0.6748 0.0234
-3.500 0.2297 0.01769 0.01378 -0.1071 0.6672 0.0239
-3.250 0.2564 0.01606 0.01204 -0.1088 0.6600 0.0249
-3.000 0.2934 0.01415 0.00988 -0.1116 0.6549 0.0272
-2.750 0.3260 0.01229 0.00781 -0.1136 0.6501 0.0274
-2.500 0.3556 0.01073 0.00606 -0.1149 0.6447 0.0275
-2.000 0.4281 0.02207 0.01672 -0.1235 0.6434 0.0278
-1.750 0.4559 0.02054 0.01506 -0.1243 0.6364 0.0281
-1.500 0.4841 0.01931 0.01368 -0.1248 0.6276 0.0284
-1.250 0.5143 0.01792 0.01204 -0.1250 0.6191 0.0271
-1.000 0.5428 0.01667 0.01060 -0.1254 0.6099 0.0267
-0.750 0.5720 0.01553 0.00925 -0.1256 0.6016 0.0265
-0.500 0.6010 0.01464 0.00811 -0.1256 0.5928 0.0270
-0.250 0.6297 0.01389 0.00718 -0.1257 0.5836 0.0271
0.000 0.6579 0.01314 0.00625 -0.1258 0.5739 0.0268
0.250 0.6860 0.01253 0.00550 -0.1257 0.5628 0.0266
0.500 0.7139 0.01206 0.00488 -0.1257 0.5492 0.0267
0.750 0.7412 0.01173 0.00439 -0.1255 0.5314 0.0269
1.000 0.7679 0.01155 0.00404 -0.1252 0.5055 0.0272
1.250 0.7937 0.01155 0.00384 -0.1248 0.4747 0.0275
1.500 0.8195 0.01139 0.00353 -0.1244 0.4514 0.0278
1.750 0.8458 0.01128 0.00334 -0.1242 0.4354 0.0284
2.000 0.8723 0.01127 0.00330 -0.1240 0.4232 0.0295
2.250 0.8990 0.01126 0.00326 -0.1239 0.4129 0.0302
2.500 0.9255 0.01128 0.00323 -0.1237 0.4015 0.0304
2.750 0.9519 0.01135 0.00324 -0.1234 0.3891 0.0308
3.000 0.9784 0.01141 0.00327 -0.1233 0.3765 0.0313
3.250 1.0050 0.01149 0.00332 -0.1231 0.3651 0.0321
3.500 1.0314 0.01159 0.00340 -0.1229 0.3563 0.0329
3.750 1.0578 0.01171 0.00349 -0.1227 0.3454 0.0336
4.000 1.0840 0.01184 0.00358 -0.1224 0.3332 0.0365
4.250 1.1097 0.01203 0.00372 -0.1221 0.3190 0.0387
4.500 1.1353 0.01224 0.00387 -0.1218 0.3048 0.0409
4.750 1.1610 0.01231 0.00416 -0.1217 0.2894 0.2307
5.250 1.2058 0.01176 0.00474 -0.1200 0.2554 1.0000
5.500 1.2299 0.01213 0.00501 -0.1195 0.2389 1.0000
5.750 1.2533 0.01255 0.00532 -0.1189 0.2206 1.0000
6.000 1.2763 0.01300 0.00568 -0.1182 0.1998 1.0000
6.250 1.2980 0.01356 0.00609 -0.1174 0.1734 1.0000
6.500 1.3180 0.01427 0.00660 -0.1163 0.1423 1.0000
6.750 1.3397 0.01479 0.00705 -0.1155 0.1293 1.0000
7.000 1.3613 0.01529 0.00750 -0.1146 0.1195 1.0000
7.250 1.3836 0.01571 0.00791 -0.1138 0.1131 1.0000
7.500 1.4051 0.01617 0.00835 -0.1130 0.1071 1.0000
7.750 1.4276 0.01653 0.00875 -0.1122 0.1035 1.0000
8.000 1.4496 0.01691 0.00914 -0.1115 0.0973 1.0000
8.250 1.4699 0.01741 0.00962 -0.1104 0.0881 1.0000
8.750 1.5020 0.01896 0.01097 -0.1071 0.0548 1.0000
9.000 1.5184 0.01963 0.01165 -0.1055 0.0499 1.0000
9.250 1.5351 0.02016 0.01223 -0.1039 0.0459 1.0000
9.500 1.5484 0.02084 0.01293 -0.1018 0.0403 1.0000
9.750 1.5556 0.02194 0.01394 -0.0989 0.0210 1.0000
10.000 1.5631 0.02307 0.01505 -0.0962 0.0159 1.0000
10.250 1.5729 0.02408 0.01613 -0.0939 0.0141 1.0000
10.500 1.5819 0.02519 0.01732 -0.0918 0.0127 1.0000
10.750 1.5931 0.02620 0.01841 -0.0900 0.0120 1.0000
11.000 1.6029 0.02734 0.01965 -0.0882 0.0114 1.0000
11.250 1.6116 0.02862 0.02102 -0.0865 0.0108 1.0000
11.500 1.6188 0.03007 0.02256 -0.0847 0.0103 1.0000
11.750 1.6242 0.03174 0.02433 -0.0830 0.0098 1.0000
12.000 1.6268 0.03372 0.02642 -0.0812 0.0094 1.0000
12.250 1.6331 0.03543 0.02822 -0.0798 0.0091 1.0000
12.500 1.6385 0.03728 0.03019 -0.0786 0.0088 1.0000
12.750 1.6424 0.03932 0.03233 -0.0774 0.0085 1.0000
13.000 1.6452 0.04157 0.03468 -0.0763 0.0082 1.0000
13.250 1.6469 0.04399 0.03721 -0.0754 0.0079 1.0000
13.500 1.6473 0.04662 0.03994 -0.0746 0.0076 1.0000
13.750 1.6462 0.04950 0.04293 -0.0739 0.0075 1.0000
14.000 1.6434 0.05268 0.04622 -0.0734 0.0073 1.0000
14.250 1.6387 0.05618 0.04984 -0.0730 0.0071 1.0000
14.500 1.6320 0.06004 0.05382 -0.0728 0.0070 1.0000
14.750 1.6227 0.06440 0.05830 -0.0729 0.0068 1.0000
15.000 1.6112 0.06921 0.06324 -0.0732 0.0067 1.0000
15.250 1.6047 0.07341 0.06757 -0.0736 0.0067 1.0000
15.500 1.5973 0.07787 0.07215 -0.0742 0.0066 1.0000
15.750 1.5889 0.08262 0.07703 -0.0751 0.0065 1.0000
16.000 1.5799 0.08753 0.08207 -0.0760 0.0064 1.0000
16.250 1.5702 0.09267 0.08733 -0.0772 0.0064 1.0000
16.500 1.5603 0.09792 0.09272 -0.0786 0.0063 1.0000
16.750 1.5502 0.10328 0.09820 -0.0801 0.0062 1.0000
17.000 1.5399 0.10868 0.10372 -0.0817 0.0061 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 81 AIRFOIL (goe81-il)