FX 63-145 AIRFOIL (fx63145-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 63-145 AIRFOIL (fx63145-il) Reynolds number: 1,000,000 Max Cl/Cd: 86.92 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx63145-il-1000000.txt Download as CSV file: xf-fx63145-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-145 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1558 0.06517 0.06033 -0.1021 0.1420 0.0089
-9.000 -0.1813 0.05476 0.04993 -0.1104 0.1617 0.0087
-8.750 -0.2070 0.04890 0.04401 -0.1142 0.1605 0.0082
-8.500 -0.2296 0.04560 0.04067 -0.1145 0.1625 0.0082
-8.250 -0.2369 0.04243 0.03738 -0.1161 0.1625 0.0086
-7.750 -0.2438 0.03281 0.02701 -0.1170 0.1616 0.0101
-7.500 -0.2266 0.03141 0.02551 -0.1171 0.1337 0.0103
-7.250 -0.2075 0.03011 0.02410 -0.1174 0.1323 0.0106
-7.000 -0.1869 0.02878 0.02264 -0.1177 0.1299 0.0110
-6.750 -0.1644 0.02728 0.02091 -0.1178 0.1277 0.0122
-6.500 -0.1424 0.02451 0.01768 -0.1182 0.1259 0.0136
-6.250 -0.1178 0.02349 0.01664 -0.1185 0.1242 0.0141
-6.000 -0.0921 0.02252 0.01554 -0.1189 0.1236 0.0149
-5.750 -0.0627 0.02314 0.01592 -0.1187 0.1222 0.0169
-5.500 -0.0381 0.01992 0.01249 -0.1194 0.1211 0.0181
-5.250 -0.0112 0.01908 0.01158 -0.1196 0.1185 0.0190
-5.000 0.0174 0.01887 0.01121 -0.1197 0.1157 0.0211
-4.750 0.0443 0.01743 0.00967 -0.1200 0.1140 0.0237
-4.500 0.0724 0.01688 0.00900 -0.1203 0.1120 0.0281
-3.750 0.1579 0.01356 0.00535 -0.1189 0.0933 0.0121
-2.750 0.2791 0.01208 0.00358 -0.1211 0.0900 0.0090
-2.500 0.3090 0.01193 0.00336 -0.1215 0.0898 0.0088
-2.250 0.3384 0.01184 0.00321 -0.1218 0.0896 0.0087
-2.000 0.3677 0.01178 0.00311 -0.1220 0.0894 0.0088
-1.750 0.3968 0.01176 0.00304 -0.1222 0.0892 0.0092
-1.500 0.4258 0.01175 0.00301 -0.1224 0.0890 0.0111
-1.250 0.4548 0.01173 0.00301 -0.1226 0.0888 0.0316
-1.000 0.4833 0.01178 0.00303 -0.1227 0.0886 0.0348
-0.750 0.5117 0.01184 0.00308 -0.1228 0.0884 0.0383
-0.500 0.5401 0.01189 0.00311 -0.1229 0.0882 0.0410
-0.250 0.5686 0.01194 0.00315 -0.1231 0.0880 0.0447
0.000 0.5969 0.01199 0.00321 -0.1232 0.0877 0.0514
0.250 0.6250 0.01207 0.00330 -0.1233 0.0875 0.0632
0.500 0.6531 0.01216 0.00338 -0.1234 0.0873 0.0714
0.750 0.6812 0.01224 0.00351 -0.1235 0.0870 0.0977
1.000 0.7092 0.01233 0.00360 -0.1236 0.0868 0.1042
1.250 0.7368 0.01246 0.00372 -0.1236 0.0865 0.1054
1.500 0.7641 0.01261 0.00387 -0.1235 0.0860 0.1063
1.750 0.7913 0.01276 0.00402 -0.1234 0.0854 0.1071
2.000 0.8184 0.01293 0.00419 -0.1234 0.0850 0.1079
2.250 0.8453 0.01310 0.00438 -0.1232 0.0845 0.1097
2.500 0.8721 0.01328 0.00457 -0.1231 0.0841 0.1125
2.750 0.8988 0.01348 0.00480 -0.1230 0.0835 0.1163
3.000 0.9251 0.01369 0.00507 -0.1229 0.0829 0.1359
3.250 0.9508 0.01397 0.00538 -0.1226 0.0823 0.1412
3.500 0.9781 0.01407 0.00550 -0.1226 0.0821 0.1477
3.750 1.0055 0.01416 0.00559 -0.1226 0.0818 0.1520
4.000 1.0329 0.01425 0.00572 -0.1226 0.0814 0.1598
4.250 1.0604 0.01434 0.00586 -0.1227 0.0810 0.1917
4.500 1.0872 0.01447 0.00600 -0.1227 0.0805 0.1992
4.750 1.1148 0.01454 0.00618 -0.1228 0.0799 0.2413
5.000 1.1419 0.01465 0.00636 -0.1229 0.0795 0.2704
5.250 1.1694 0.01472 0.00650 -0.1231 0.0784 0.3156
5.500 1.1981 0.01469 0.00660 -0.1235 0.0765 0.3789
5.750 1.2264 0.01470 0.00678 -0.1240 0.0754 0.4561
6.000 1.2548 0.01470 0.00695 -0.1244 0.0743 0.5333
6.250 1.2821 0.01475 0.00708 -0.1245 0.0722 0.5675
6.500 1.3066 0.01504 0.00727 -0.1241 0.0585 0.5818
6.750 1.3308 0.01533 0.00760 -0.1237 0.0541 0.5963
7.000 1.3547 0.01565 0.00800 -0.1231 0.0526 0.6124
7.250 1.3771 0.01606 0.00848 -0.1224 0.0500 0.6237
7.500 1.3986 0.01651 0.00897 -0.1215 0.0464 0.6314
7.750 1.4229 0.01674 0.00923 -0.1211 0.0447 0.6385
8.000 1.4416 0.01734 0.00975 -0.1198 0.0261 0.6459
8.250 1.4548 0.01827 0.01061 -0.1177 0.0096 0.6529
8.500 1.4708 0.01895 0.01129 -0.1160 0.0034 0.6621
8.750 1.4895 0.01940 0.01180 -0.1147 0.0028 0.6719
9.000 1.5045 0.01989 0.01234 -0.1127 0.0027 0.6822
9.250 1.5185 0.02041 0.01295 -0.1106 0.0026 0.6917
9.500 1.5316 0.02099 0.01360 -0.1085 0.0026 0.7015
9.750 1.5434 0.02165 0.01435 -0.1062 0.0025 0.7131
10.000 1.5536 0.02241 0.01522 -0.1039 0.0024 0.7273
10.250 1.5623 0.02330 0.01621 -0.1015 0.0023 0.7438
10.500 1.5701 0.02428 0.01731 -0.0991 0.0022 0.7615
10.750 1.5772 0.02536 0.01854 -0.0970 0.0022 0.7929
11.000 1.5822 0.02656 0.01995 -0.0948 0.0021 0.8483
11.250 1.5795 0.02775 0.02141 -0.0914 0.0021 1.0000
11.500 1.5828 0.02950 0.02324 -0.0897 0.0021 1.0000
11.750 1.5850 0.03149 0.02533 -0.0884 0.0020 1.0000
12.000 1.5856 0.03383 0.02775 -0.0873 0.0020 1.0000
12.250 1.5849 0.03644 0.03047 -0.0865 0.0020 1.0000
12.500 1.5818 0.03949 0.03363 -0.0859 0.0019 1.0000
12.750 1.5805 0.04250 0.03673 -0.0856 0.0019 1.0000
13.000 1.5808 0.04542 0.03974 -0.0854 0.0019 1.0000
13.250 1.5795 0.04862 0.04303 -0.0854 0.0018 1.0000
13.500 1.5766 0.05209 0.04660 -0.0856 0.0018 1.0000
13.750 1.5729 0.05574 0.05035 -0.0859 0.0018 1.0000
14.000 1.5678 0.05972 0.05443 -0.0864 0.0018 1.0000
14.250 1.5617 0.06394 0.05876 -0.0871 0.0017 1.0000
14.500 1.5551 0.06839 0.06332 -0.0879 0.0017 1.0000
14.750 1.5475 0.07317 0.06821 -0.0891 0.0017 1.0000
15.000 1.5387 0.07828 0.07344 -0.0904 0.0016 1.0000
15.250 1.5308 0.08338 0.07866 -0.0919 0.0016 1.0000
15.500 1.5223 0.08876 0.08414 -0.0936 0.0016 1.0000
15.750 1.5122 0.09454 0.09004 -0.0956 0.0016 1.0000
16.000 1.5024 0.10042 0.09604 -0.0978 0.0015 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 63-145 AIRFOIL (fx63145-il)