FX 71-L-150/20 AIRFOIL (fx711520-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 71-L-150/20 AIRFOIL (fx711520-il) Reynolds number: 50,000 Max Cl/Cd: 27.18 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx711520-il-50000.txt Download as CSV file: xf-fx711520-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 71-L-150/20 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6790 0.10377 0.09656 -0.0093 1.0000 0.1221
-11.500 -0.7944 0.08350 0.07613 -0.0246 1.0000 0.1121
-11.250 -0.8144 0.07766 0.07022 -0.0267 1.0000 0.1108
-11.000 -0.8428 0.07209 0.06450 -0.0284 1.0000 0.1096
-10.750 -0.8695 0.06748 0.05971 -0.0286 1.0000 0.1083
-10.500 -0.8926 0.06344 0.05542 -0.0275 1.0000 0.1072
-10.250 -0.9045 0.05944 0.05111 -0.0265 1.0000 0.1066
-10.000 -0.9075 0.05572 0.04705 -0.0255 1.0000 0.1064
-9.750 -0.9027 0.05225 0.04326 -0.0247 1.0000 0.1070
-9.500 -0.8921 0.04903 0.03974 -0.0240 1.0000 0.1082
-9.250 -0.8782 0.04597 0.03632 -0.0233 1.0000 0.1104
-9.000 -0.8623 0.04308 0.03300 -0.0225 1.0000 0.1134
-8.750 -0.8327 0.04060 0.03066 -0.0229 1.0000 0.1205
-8.500 -0.8058 0.03822 0.02829 -0.0228 1.0000 0.1304
-8.250 -0.7774 0.03598 0.02613 -0.0227 1.0000 0.1463
-8.000 -0.7559 0.03397 0.02440 -0.0221 1.0000 0.1707
-7.750 -0.7431 0.03214 0.02292 -0.0209 1.0000 0.2024
-7.500 -0.7329 0.03017 0.02137 -0.0191 1.0000 0.2438
-7.250 -0.7309 0.02831 0.02019 -0.0160 1.0000 0.2962
-7.000 -0.7387 0.02717 0.01977 -0.0110 1.0000 0.3579
-6.750 -0.7426 0.02782 0.02111 -0.0045 1.0000 0.4308
-6.500 -0.7320 0.03064 0.02413 0.0021 1.0000 0.4905
-6.250 -0.7039 0.03481 0.02821 0.0080 1.0000 0.5290
-6.000 -0.6478 0.04075 0.03392 0.0131 1.0000 0.5591
-5.750 -0.6469 0.04160 0.03467 0.0183 1.0000 0.5830
-5.500 -0.6196 0.04387 0.03678 0.0226 1.0000 0.6064
-5.250 -0.5450 0.04743 0.04001 0.0234 1.0000 0.6340
-5.000 -0.5222 0.04797 0.04040 0.0258 1.0000 0.6569
-4.750 -0.5051 0.04802 0.04034 0.0282 1.0000 0.6784
-4.500 -0.4879 0.04783 0.04003 0.0301 1.0000 0.6992
-4.250 -0.4708 0.04749 0.03956 0.0318 1.0000 0.7194
-4.000 -0.4546 0.04698 0.03894 0.0332 1.0000 0.7385
-3.750 -0.4464 0.04631 0.03819 0.0353 1.0000 0.7556
-3.500 -0.4870 0.04513 0.03704 0.0435 1.0000 0.7687
-3.250 -0.4656 0.04449 0.03628 0.0438 1.0000 0.7842
-3.000 -0.4384 0.04386 0.03552 0.0431 1.0000 0.7999
-2.750 -0.4128 0.04319 0.03475 0.0423 0.9995 0.8151
-2.500 -0.3589 0.04266 0.03405 0.0362 0.9932 0.8304
-2.250 -0.3415 0.04192 0.03320 0.0353 0.9858 0.8423
-2.000 -0.2800 0.04150 0.03264 0.0278 0.9796 0.8552
-1.750 -0.2331 0.04098 0.03202 0.0227 0.9729 0.8670
-1.500 -0.2104 0.04061 0.03157 0.0210 0.9663 0.8776
-1.250 -0.1591 0.04014 0.03103 0.0150 0.9594 0.8865
-1.000 -0.1277 0.03989 0.03072 0.0119 0.9530 0.8951
-0.750 -0.1042 0.03957 0.03037 0.0104 0.9469 0.9023
-0.500 -0.0669 0.03941 0.03018 0.0065 0.9405 0.9109
-0.250 -0.0314 0.03938 0.03013 0.0029 0.9341 0.9194
0.000 0.0001 0.03927 0.03002 0.0000 0.9266 0.9267
0.250 0.0321 0.03938 0.03013 -0.0030 0.9193 0.9342
0.500 0.0666 0.03941 0.03018 -0.0065 0.9110 0.9405
0.750 0.1037 0.03956 0.03036 -0.0103 0.9024 0.9469
1.000 0.1279 0.03988 0.03071 -0.0120 0.8951 0.9530
1.250 0.1597 0.04013 0.03102 -0.0151 0.8866 0.9595
1.500 0.2110 0.04061 0.03157 -0.0211 0.8777 0.9664
1.750 0.2323 0.04097 0.03201 -0.0226 0.8673 0.9729
2.000 0.2800 0.04149 0.03263 -0.0278 0.8553 0.9795
2.250 0.3431 0.04189 0.03318 -0.0356 0.8422 0.9859
2.500 0.3593 0.04264 0.03403 -0.0363 0.8304 0.9932
2.750 0.4131 0.04317 0.03472 -0.0423 0.8149 0.9995
3.000 0.4387 0.04384 0.03550 -0.0431 0.7999 1.0000
3.250 0.4670 0.04447 0.03626 -0.0440 0.7844 1.0000
3.500 0.4802 0.04520 0.03709 -0.0427 0.7693 1.0000
3.750 0.4461 0.04629 0.03817 -0.0353 0.7557 1.0000
4.000 0.4544 0.04697 0.03894 -0.0332 0.7386 1.0000
4.250 0.4711 0.04747 0.03954 -0.0318 0.7194 1.0000
4.500 0.4883 0.04781 0.03999 -0.0302 0.6991 1.0000
4.750 0.5049 0.04801 0.04032 -0.0282 0.6785 1.0000
5.000 0.5223 0.04795 0.04038 -0.0258 0.6570 1.0000
5.250 0.5445 0.04743 0.04001 -0.0233 0.6342 1.0000
5.500 0.6193 0.04387 0.03678 -0.0225 0.6064 1.0000
5.750 0.6472 0.04156 0.03463 -0.0183 0.5830 1.0000
6.000 0.6479 0.04073 0.03391 -0.0131 0.5591 1.0000
6.250 0.7035 0.03484 0.02824 -0.0080 0.5291 1.0000
6.500 0.7325 0.03056 0.02405 -0.0020 0.4903 1.0000
6.750 0.7425 0.02781 0.02110 0.0045 0.4309 1.0000
7.000 0.7386 0.02717 0.01977 0.0110 0.3579 1.0000
7.250 0.7309 0.02831 0.02019 0.0160 0.2961 1.0000
7.500 0.7327 0.03017 0.02138 0.0191 0.2434 1.0000
7.750 0.7425 0.03214 0.02291 0.0209 0.2019 1.0000
8.000 0.7560 0.03397 0.02440 0.0221 0.1707 1.0000
8.250 0.7775 0.03598 0.02613 0.0227 0.1463 1.0000
8.500 0.8057 0.03822 0.02829 0.0228 0.1301 1.0000
8.750 0.8326 0.04061 0.03067 0.0229 0.1203 1.0000
9.000 0.8629 0.04311 0.03302 0.0225 0.1133 1.0000
9.250 0.8783 0.04598 0.03634 0.0233 0.1103 1.0000
9.500 0.8923 0.04904 0.03975 0.0239 0.1082 1.0000
9.750 0.9029 0.05226 0.04327 0.0247 0.1070 1.0000
10.000 0.9078 0.05574 0.04707 0.0255 0.1063 1.0000
10.250 0.9048 0.05946 0.05113 0.0265 0.1066 1.0000
10.500 0.8927 0.06348 0.05546 0.0275 0.1072 1.0000
10.750 0.8700 0.06749 0.05972 0.0286 0.1083 1.0000
11.000 0.8431 0.07213 0.06454 0.0283 0.1096 1.0000
11.250 0.8151 0.07768 0.07023 0.0266 0.1108 1.0000
11.500 0.7960 0.08347 0.07610 0.0245 0.1122 1.0000
11.750 0.6791 0.10396 0.09676 0.0090 0.1221 1.0000
12.000 0.6820 0.10955 0.10232 0.0073 0.1242 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 71-L-150/20 AIRFOIL (fx711520-il)