FX 71-L-150/25 AIRFOIL (fx711525-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 71-L-150/25 AIRFOIL (fx711525-il) Reynolds number: 50,000 Max Cl/Cd: 27.36 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx711525-il-50000.txt Download as CSV file: xf-fx711525-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 71-L-150/25 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6748 0.10616 0.09880 -0.0140 1.0000 0.1242
-11.750 -0.6784 0.09969 0.09235 -0.0165 1.0000 0.1219
-11.500 -0.8141 0.08008 0.07248 -0.0311 1.0000 0.1123
-11.250 -0.8303 0.07498 0.06729 -0.0322 1.0000 0.1114
-11.000 -0.8510 0.07031 0.06249 -0.0326 1.0000 0.1104
-10.750 -0.8726 0.06624 0.05825 -0.0317 1.0000 0.1095
-10.500 -0.8898 0.06245 0.05422 -0.0302 1.0000 0.1087
-10.250 -0.8989 0.05866 0.05014 -0.0288 1.0000 0.1084
-10.000 -0.9012 0.05509 0.04626 -0.0275 1.0000 0.1083
-9.750 -0.8984 0.05169 0.04250 -0.0262 1.0000 0.1090
-9.500 -0.8906 0.04848 0.03892 -0.0249 1.0000 0.1103
-9.250 -0.8801 0.04551 0.03547 -0.0236 1.0000 0.1122
-9.000 -0.8540 0.04261 0.03261 -0.0237 1.0000 0.1169
-8.750 -0.8341 0.04024 0.02998 -0.0229 1.0000 0.1238
-8.500 -0.8069 0.03788 0.02773 -0.0229 1.0000 0.1350
-8.250 -0.7823 0.03566 0.02550 -0.0223 1.0000 0.1516
-8.000 -0.7637 0.03378 0.02405 -0.0215 1.0000 0.1769
-7.750 -0.7510 0.03198 0.02258 -0.0199 1.0000 0.2080
-7.500 -0.7408 0.03005 0.02104 -0.0177 1.0000 0.2487
-7.250 -0.7375 0.02828 0.01992 -0.0144 1.0000 0.2992
-7.000 -0.7436 0.02717 0.01951 -0.0092 1.0000 0.3588
-6.750 -0.7475 0.02747 0.02047 -0.0029 1.0000 0.4273
-6.500 -0.7385 0.02966 0.02292 0.0035 1.0000 0.4882
-6.250 -0.7229 0.03223 0.02545 0.0092 1.0000 0.5299
-6.000 -0.6745 0.03723 0.03031 0.0140 1.0000 0.5628
-5.750 -0.6746 0.03796 0.03092 0.0193 1.0000 0.5883
-5.500 -0.5812 0.04412 0.03674 0.0208 1.0000 0.6160
-5.250 -0.5619 0.04510 0.03757 0.0243 1.0000 0.6393
-5.000 -0.5514 0.04544 0.03781 0.0279 1.0000 0.6609
-4.750 -0.5420 0.04553 0.03779 0.0313 1.0000 0.6817
-4.500 -0.5306 0.04548 0.03764 0.0342 1.0000 0.7017
-4.250 -0.5202 0.04523 0.03730 0.0370 1.0000 0.7213
-4.000 -0.5100 0.04484 0.03680 0.0396 1.0000 0.7400
-3.750 -0.5049 0.04425 0.03613 0.0425 1.0000 0.7578
-3.500 -0.4800 0.04383 0.03559 0.0428 1.0000 0.7752
-3.250 -0.4700 0.04311 0.03478 0.0446 1.0000 0.7910
-3.000 -0.4646 0.04236 0.03395 0.0469 1.0000 0.8063
-2.750 -0.4211 0.04200 0.03343 0.0437 1.0000 0.8222
-2.500 -0.3869 0.04148 0.03280 0.0415 1.0000 0.8358
-2.250 -0.3826 0.04070 0.03197 0.0434 1.0000 0.8478
-2.000 -0.3786 0.04002 0.03122 0.0454 1.0000 0.8600
-1.750 -0.3038 0.03998 0.03102 0.0356 0.9932 0.8729
-1.500 -0.2696 0.03960 0.03056 0.0320 0.9856 0.8844
-1.250 -0.2116 0.03943 0.03028 0.0246 0.9785 0.8951
-1.000 -0.1732 0.03917 0.02995 0.0203 0.9708 0.9045
-0.750 -0.1271 0.03901 0.02974 0.0148 0.9637 0.9142
-0.500 -0.0865 0.03889 0.02958 0.0102 0.9562 0.9231
-0.250 -0.0419 0.03882 0.02949 0.0049 0.9486 0.9324
0.000 -0.0002 0.03880 0.02945 0.0000 0.9405 0.9404
0.250 0.0420 0.03882 0.02948 -0.0049 0.9324 0.9486
0.500 0.0868 0.03888 0.02957 -0.0102 0.9232 0.9563
0.750 0.1269 0.03901 0.02973 -0.0148 0.9143 0.9637
1.000 0.1732 0.03917 0.02995 -0.0203 0.9045 0.9709
1.250 0.2116 0.03942 0.03027 -0.0246 0.8952 0.9785
1.500 0.2691 0.03959 0.03054 -0.0320 0.8844 0.9856
1.750 0.3036 0.03997 0.03101 -0.0355 0.8728 0.9931
2.000 0.3792 0.04000 0.03120 -0.0455 0.8602 1.0000
2.250 0.3823 0.04069 0.03196 -0.0434 0.8479 1.0000
2.500 0.3868 0.04146 0.03279 -0.0415 0.8359 1.0000
2.750 0.4213 0.04198 0.03342 -0.0438 0.8222 1.0000
3.000 0.4650 0.04233 0.03393 -0.0470 0.8063 1.0000
3.250 0.4690 0.04310 0.03477 -0.0445 0.7911 1.0000
3.500 0.4781 0.04384 0.03559 -0.0426 0.7754 1.0000
3.750 0.5054 0.04423 0.03611 -0.0426 0.7578 1.0000
4.000 0.5113 0.04481 0.03677 -0.0397 0.7401 1.0000
4.250 0.5206 0.04521 0.03728 -0.0370 0.7213 1.0000
4.500 0.5310 0.04544 0.03761 -0.0343 0.7018 1.0000
4.750 0.5425 0.04551 0.03778 -0.0313 0.6818 1.0000
5.000 0.5519 0.04542 0.03778 -0.0279 0.6610 1.0000
5.250 0.5620 0.04509 0.03755 -0.0243 0.6394 1.0000
5.500 0.5808 0.04413 0.03675 -0.0208 0.6161 1.0000
5.750 0.6734 0.03805 0.03101 -0.0193 0.5884 1.0000
6.000 0.6744 0.03723 0.03030 -0.0140 0.5628 1.0000
6.250 0.7228 0.03223 0.02544 -0.0092 0.5299 1.0000
6.500 0.7384 0.02966 0.02292 -0.0035 0.4883 1.0000
6.750 0.7474 0.02747 0.02047 0.0029 0.4275 1.0000
7.000 0.7434 0.02717 0.01950 0.0093 0.3586 1.0000
7.250 0.7373 0.02828 0.01991 0.0144 0.2991 1.0000
7.500 0.7409 0.03005 0.02103 0.0177 0.2487 1.0000
7.750 0.7512 0.03198 0.02258 0.0199 0.2080 1.0000
8.000 0.7634 0.03379 0.02406 0.0215 0.1765 1.0000
8.250 0.7825 0.03566 0.02550 0.0223 0.1516 1.0000
8.500 0.8071 0.03788 0.02772 0.0228 0.1351 1.0000
8.750 0.8343 0.04026 0.03000 0.0229 0.1237 1.0000
9.000 0.8542 0.04261 0.03261 0.0236 0.1168 1.0000
9.250 0.8803 0.04551 0.03547 0.0236 0.1122 1.0000
9.500 0.8907 0.04848 0.03891 0.0249 0.1103 1.0000
9.750 0.8984 0.05170 0.04251 0.0261 0.1090 1.0000
10.000 0.9016 0.05508 0.04624 0.0274 0.1084 1.0000
10.250 0.8989 0.05869 0.05017 0.0288 0.1083 1.0000
10.500 0.8904 0.06244 0.05421 0.0301 0.1088 1.0000
10.750 0.8725 0.06628 0.05828 0.0317 0.1094 1.0000
11.000 0.8522 0.07029 0.06247 0.0325 0.1104 1.0000
11.250 0.8301 0.07503 0.06735 0.0321 0.1113 1.0000
11.500 0.8155 0.08008 0.07248 0.0310 0.1123 1.0000
11.750 0.6795 0.09966 0.09231 0.0165 0.1220 1.0000
12.000 0.6766 0.10600 0.09864 0.0141 0.1242 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 71-L-150/25 AIRFOIL (fx711525-il)