GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il) Reynolds number: 500,000 Max Cl/Cd: 92.87 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe198-il-500000-n5.txt Download as CSV file: xf-goe198-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 198 (L.F.G. 5294) 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2017 0.09874 0.09598 -0.0489 0.8779 0.0197
-8.500 -0.1944 0.09573 0.09297 -0.0512 0.8737 0.0198
-8.250 -0.1886 0.09272 0.08994 -0.0541 0.8693 0.0199
-8.000 -0.1784 0.08971 0.08691 -0.0542 0.8662 0.0200
-7.750 -0.1680 0.08717 0.08438 -0.0550 0.8619 0.0201
-7.500 -0.1583 0.08483 0.08204 -0.0561 0.8575 0.0204
-7.250 -0.1453 0.08230 0.07949 -0.0583 0.8538 0.0207
-7.000 -0.1309 0.07967 0.07684 -0.0609 0.8502 0.0211
-6.750 -0.1147 0.07686 0.07402 -0.0642 0.8458 0.0217
-6.500 -0.0951 0.07367 0.07081 -0.0690 0.8411 0.0230
-6.250 -0.0681 0.06974 0.06680 -0.0775 0.8365 0.0234
-6.000 -0.0420 0.06598 0.06300 -0.0840 0.8314 0.0235
-5.750 -0.0151 0.06230 0.05926 -0.0897 0.8259 0.0235
-5.500 0.0122 0.05874 0.05562 -0.0946 0.8211 0.0236
-5.250 0.0320 0.05564 0.05251 -0.0966 0.8161 0.0237
-5.000 0.0526 0.05336 0.05020 -0.0979 0.8103 0.0240
-4.750 0.0767 0.05111 0.04791 -0.1002 0.8052 0.0243
-4.500 0.1034 0.04890 0.04566 -0.1030 0.8002 0.0250
-4.250 0.1336 0.04639 0.04308 -0.1067 0.7944 0.0261
-4.000 0.1823 0.04227 0.03875 -0.1151 0.7889 0.0278
-3.750 0.2182 0.03905 0.03539 -0.1194 0.7821 0.0279
-3.500 0.2517 0.03619 0.03239 -0.1226 0.7744 0.0279
-3.250 0.2849 0.03349 0.02955 -0.1255 0.7675 0.0280
-3.000 0.3145 0.03074 0.02669 -0.1280 0.7601 0.0284
-2.750 0.3418 0.02946 0.02535 -0.1292 0.7524 0.0289
-2.500 0.3774 0.02682 0.02252 -0.1317 0.7437 0.0273
-2.250 0.4097 0.02475 0.02027 -0.1336 0.7330 0.0264
-2.000 0.4450 0.02220 0.01745 -0.1357 0.7196 0.0259
-1.750 0.4799 0.01977 0.01470 -0.1374 0.7043 0.0259
-1.500 0.5143 0.01744 0.01200 -0.1389 0.6883 0.0259
-1.250 0.5479 0.01523 0.00936 -0.1400 0.6711 0.0266
-1.000 0.5803 0.01322 0.00686 -0.1410 0.6517 0.0266
-0.750 0.6097 0.01239 0.00568 -0.1413 0.6285 0.0268
-0.500 0.6381 0.01196 0.00497 -0.1413 0.6022 0.0272
-0.250 0.6658 0.01177 0.00451 -0.1413 0.5752 0.0275
0.000 0.6934 0.01167 0.00421 -0.1413 0.5524 0.0278
0.250 0.7214 0.01149 0.00391 -0.1414 0.5369 0.0282
0.500 0.7496 0.01129 0.00362 -0.1415 0.5254 0.0287
0.750 0.7776 0.01121 0.00350 -0.1416 0.5135 0.0293
1.000 0.8056 0.01119 0.00344 -0.1416 0.5005 0.0303
1.250 0.8334 0.01118 0.00338 -0.1417 0.4866 0.0312
1.500 0.8611 0.01120 0.00333 -0.1417 0.4686 0.0318
1.750 0.8886 0.01127 0.00331 -0.1417 0.4477 0.0325
2.000 0.9156 0.01141 0.00333 -0.1416 0.4189 0.0332
2.250 0.9411 0.01171 0.00341 -0.1413 0.3751 0.0344
2.500 0.9669 0.01200 0.00355 -0.1411 0.3473 0.0361
2.750 0.9932 0.01223 0.00370 -0.1409 0.3308 0.0385
3.250 1.0463 0.01263 0.00408 -0.1406 0.3068 0.0704
3.500 1.0726 0.01284 0.00426 -0.1404 0.2960 0.0800
3.750 1.0987 0.01307 0.00445 -0.1402 0.2841 0.0877
4.000 1.1250 0.01325 0.00462 -0.1401 0.2723 0.0998
4.250 1.1507 0.01239 0.00509 -0.1404 0.2579 0.8391
4.750 1.1890 0.01352 0.00576 -0.1374 0.1619 1.0000
5.000 1.2040 0.01495 0.00669 -0.1355 0.0708 1.0000
5.250 1.2275 0.01540 0.00707 -0.1349 0.0622 1.0000
5.500 1.2518 0.01575 0.00739 -0.1344 0.0599 1.0000
5.750 1.2758 0.01610 0.00774 -0.1338 0.0582 1.0000
6.000 1.2995 0.01647 0.00811 -0.1332 0.0568 1.0000
6.250 1.3230 0.01684 0.00849 -0.1325 0.0557 1.0000
6.500 1.3459 0.01723 0.00890 -0.1318 0.0548 1.0000
6.750 1.3685 0.01764 0.00933 -0.1310 0.0541 1.0000
7.000 1.3908 0.01803 0.00976 -0.1302 0.0538 1.0000
7.250 1.4127 0.01844 0.01019 -0.1293 0.0532 1.0000
7.500 1.4340 0.01886 0.01065 -0.1284 0.0522 1.0000
7.750 1.4543 0.01931 0.01114 -0.1272 0.0511 1.0000
8.000 1.4738 0.01980 0.01167 -0.1260 0.0502 1.0000
8.250 1.4922 0.02033 0.01223 -0.1246 0.0493 1.0000
8.500 1.5094 0.02090 0.01284 -0.1230 0.0484 1.0000
8.750 1.5236 0.02154 0.01353 -0.1208 0.0474 1.0000
9.000 1.5352 0.02234 0.01437 -0.1184 0.0462 1.0000
9.250 1.5496 0.02300 0.01508 -0.1164 0.0452 1.0000
9.500 1.5662 0.02355 0.01569 -0.1149 0.0444 1.0000
9.750 1.5817 0.02419 0.01640 -0.1133 0.0436 1.0000
10.000 1.5971 0.02487 0.01713 -0.1118 0.0425 1.0000
10.250 1.6125 0.02557 0.01788 -0.1104 0.0411 1.0000
10.500 1.6269 0.02636 0.01870 -0.1089 0.0396 1.0000
10.750 1.6400 0.02726 0.01964 -0.1073 0.0381 1.0000
11.000 1.6565 0.02791 0.02038 -0.1061 0.0360 1.0000
11.250 1.6720 0.02866 0.02114 -0.1049 0.0326 1.0000
11.500 1.6857 0.02955 0.02202 -0.1036 0.0238 1.0000
11.750 1.6909 0.03116 0.02353 -0.1014 0.0177 1.0000
12.000 1.6955 0.03287 0.02530 -0.0993 0.0158 1.0000
12.250 1.7011 0.03455 0.02705 -0.0974 0.0147 1.0000
12.500 1.7065 0.03627 0.02886 -0.0957 0.0139 1.0000
12.750 1.7105 0.03817 0.03085 -0.0939 0.0132 1.0000
13.000 1.7132 0.04026 0.03303 -0.0922 0.0127 1.0000
13.250 1.7143 0.04256 0.03543 -0.0906 0.0122 1.0000
13.500 1.7156 0.04491 0.03788 -0.0891 0.0118 1.0000
13.750 1.7176 0.04726 0.04035 -0.0879 0.0115 1.0000
14.000 1.7184 0.04983 0.04303 -0.0868 0.0113 1.0000
14.250 1.7181 0.05262 0.04594 -0.0858 0.0110 1.0000
14.500 1.7167 0.05565 0.04908 -0.0850 0.0107 1.0000
14.750 1.7145 0.05892 0.05247 -0.0845 0.0105 1.0000
15.000 1.7112 0.06246 0.05612 -0.0841 0.0102 1.0000
15.250 1.7066 0.06627 0.06005 -0.0840 0.0100 1.0000
15.500 1.7005 0.07040 0.06430 -0.0840 0.0098 1.0000
15.750 1.6931 0.07485 0.06886 -0.0843 0.0096 1.0000
16.000 1.6837 0.07968 0.07381 -0.0849 0.0094 1.0000
16.250 1.6718 0.08494 0.07920 -0.0856 0.0092 1.0000
16.500 1.6621 0.08994 0.08432 -0.0864 0.0091 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il)