FX 60-160 AIRFOIL (fx60160-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 60-160 AIRFOIL (fx60160-il) Reynolds number: 500,000 Max Cl/Cd: 96.91 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx60160-il-500000.txt Download as CSV file: xf-fx60160-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-160 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.000 -0.5292 0.08378 0.08138 -0.0698 1.0000 0.0247
-12.750 -0.5686 0.07431 0.07185 -0.0744 1.0000 0.0244
-12.500 -0.5963 0.06851 0.06600 -0.0766 1.0000 0.0240
-12.250 -0.6337 0.06150 0.05886 -0.0799 0.9995 0.0239
-12.000 -0.6700 0.05099 0.04798 -0.0899 0.9951 0.0238
-11.750 -0.7018 0.04242 0.03887 -0.0969 0.9878 0.0242
-11.250 -0.6980 0.03420 0.02996 -0.1033 0.9742 0.0254
-11.000 -0.6685 0.03319 0.02892 -0.1060 0.9717 0.0260
-10.750 -0.6362 0.03234 0.02803 -0.1089 0.9700 0.0266
-10.500 -0.6161 0.03121 0.02680 -0.1094 0.9636 0.0274
-10.250 -0.5915 0.02909 0.02440 -0.1113 0.9598 0.0286
-10.000 -0.5618 0.02732 0.02227 -0.1137 0.9574 0.0297
-9.750 -0.5335 0.02436 0.01906 -0.1161 0.9554 0.0307
-9.500 -0.5126 0.02366 0.01835 -0.1157 0.9471 0.0314
-9.250 -0.4812 0.02292 0.01757 -0.1173 0.9428 0.0324
-9.000 -0.4538 0.02206 0.01660 -0.1180 0.9369 0.0335
-8.750 -0.4297 0.02101 0.01541 -0.1179 0.9283 0.0344
-8.500 -0.4039 0.02011 0.01436 -0.1180 0.9202 0.0352
-8.250 -0.3778 0.01965 0.01374 -0.1180 0.9108 0.0360
-8.000 -0.3549 0.01788 0.01190 -0.1177 0.9001 0.0373
-7.750 -0.3287 0.01717 0.01115 -0.1178 0.8904 0.0382
-7.500 -0.3027 0.01659 0.01050 -0.1178 0.8800 0.0393
-7.250 -0.2782 0.01605 0.00988 -0.1175 0.8686 0.0403
-7.000 -0.2529 0.01562 0.00937 -0.1173 0.8583 0.0417
-6.750 -0.2272 0.01523 0.00886 -0.1171 0.8490 0.0428
-6.500 -0.2029 0.01489 0.00843 -0.1167 0.8395 0.0435
-6.250 -0.1818 0.01393 0.00740 -0.1159 0.8313 0.0450
-6.000 -0.1597 0.01343 0.00688 -0.1152 0.8228 0.0464
-5.750 -0.1349 0.01311 0.00650 -0.1150 0.8160 0.0480
-5.500 -0.1111 0.01282 0.00618 -0.1145 0.8088 0.0500
-5.250 -0.0859 0.01254 0.00581 -0.1142 0.8026 0.0517
-5.000 -0.0607 0.01232 0.00554 -0.1140 0.7964 0.0529
-4.750 -0.0390 0.01168 0.00486 -0.1133 0.7903 0.0561
-4.500 -0.0129 0.01143 0.00455 -0.1133 0.7851 0.0592
-4.250 0.0123 0.01121 0.00430 -0.1130 0.7794 0.0625
-4.000 0.0377 0.01091 0.00395 -0.1128 0.7741 0.0669
-3.750 0.0643 0.01067 0.00367 -0.1128 0.7693 0.0733
-3.500 0.0897 0.01040 0.00343 -0.1126 0.7641 0.0849
-3.250 0.1095 0.00929 0.00297 -0.1121 0.7589 0.2557
-3.000 0.1299 0.00827 0.00267 -0.1116 0.7542 0.4672
-2.750 0.1552 0.00809 0.00271 -0.1112 0.7494 0.5372
-2.500 0.1819 0.00808 0.00277 -0.1110 0.7442 0.5720
-2.250 0.2099 0.00815 0.00280 -0.1110 0.7395 0.5914
-2.000 0.2380 0.00826 0.00287 -0.1110 0.7350 0.6066
-1.750 0.2652 0.00835 0.00297 -0.1108 0.7300 0.6191
-1.500 0.2930 0.00844 0.00302 -0.1108 0.7253 0.6287
-1.250 0.3218 0.00859 0.00310 -0.1110 0.7210 0.6375
-1.000 0.3488 0.00867 0.00319 -0.1108 0.7163 0.6450
-0.750 0.3764 0.00876 0.00328 -0.1107 0.7114 0.6517
-0.500 0.4044 0.00888 0.00334 -0.1107 0.7070 0.6593
-0.250 0.4321 0.00900 0.00348 -0.1106 0.7026 0.6653
0.000 0.4593 0.00904 0.00351 -0.1105 0.6975 0.6699
0.250 0.4872 0.00909 0.00350 -0.1106 0.6929 0.6738
0.500 0.5157 0.00915 0.00354 -0.1107 0.6885 0.6770
0.750 0.5421 0.00918 0.00362 -0.1105 0.6834 0.6809
1.000 0.5693 0.00924 0.00367 -0.1104 0.6784 0.6854
1.250 0.5978 0.00933 0.00369 -0.1106 0.6738 0.6897
1.500 0.6242 0.00934 0.00375 -0.1103 0.6688 0.6929
1.750 0.6509 0.00937 0.00382 -0.1101 0.6634 0.6963
2.000 0.6786 0.00944 0.00386 -0.1102 0.6586 0.7000
2.250 0.7055 0.00951 0.00393 -0.1101 0.6534 0.7039
2.500 0.7319 0.00952 0.00396 -0.1099 0.6476 0.7069
2.750 0.7591 0.00957 0.00399 -0.1098 0.6418 0.7093
3.000 0.7838 0.00957 0.00405 -0.1092 0.6338 0.7120
3.250 0.8098 0.00963 0.00407 -0.1089 0.6259 0.7150
3.500 0.8340 0.00964 0.00412 -0.1082 0.6163 0.7182
3.750 0.8580 0.00969 0.00414 -0.1075 0.6045 0.7211
4.000 0.8815 0.00972 0.00416 -0.1066 0.5933 0.7235
4.250 0.9052 0.00977 0.00424 -0.1059 0.5822 0.7260
4.500 0.9287 0.00984 0.00434 -0.1051 0.5708 0.7289
4.750 0.9514 0.00995 0.00443 -0.1041 0.5580 0.7319
5.000 0.9735 0.01008 0.00454 -0.1031 0.5441 0.7347
5.250 0.9933 0.01025 0.00465 -0.1016 0.5244 0.7374
5.500 1.0106 0.01045 0.00479 -0.0996 0.5022 0.7401
5.750 1.0274 0.01069 0.00498 -0.0975 0.4782 0.7428
6.000 1.0414 0.01101 0.00521 -0.0949 0.4543 0.7456
6.250 1.0537 0.01133 0.00544 -0.0920 0.4282 0.7487
6.500 1.0632 0.01179 0.00577 -0.0887 0.3980 0.7518
6.750 1.0689 0.01240 0.00620 -0.0847 0.3592 0.7546
7.000 1.0672 0.01335 0.00687 -0.0796 0.2998 0.7574
7.250 1.0509 0.01503 0.00799 -0.0725 0.2012 0.7607
7.500 1.0317 0.01704 0.00943 -0.0655 0.1020 0.7643
7.750 1.0351 0.01812 0.01038 -0.0621 0.0788 0.7675
8.000 1.0438 0.01894 0.01121 -0.0595 0.0701 0.7705
8.250 1.0561 0.01961 0.01194 -0.0575 0.0657 0.7736
8.500 1.0653 0.02048 0.01283 -0.0551 0.0618 0.7773
8.750 1.0738 0.02144 0.01384 -0.0528 0.0588 0.7810
9.000 1.0866 0.02218 0.01465 -0.0512 0.0571 0.7843
9.250 1.0980 0.02300 0.01555 -0.0494 0.0554 0.7880
9.500 1.1087 0.02391 0.01652 -0.0476 0.0541 0.7922
9.750 1.1185 0.02492 0.01759 -0.0459 0.0527 0.7967
10.000 1.1250 0.02618 0.01889 -0.0439 0.0507 0.8010
10.250 1.1257 0.02786 0.02066 -0.0415 0.0491 0.8058
10.500 1.1385 0.02879 0.02167 -0.0404 0.0485 0.8114
10.750 1.1500 0.02981 0.02278 -0.0392 0.0476 0.8176
11.000 1.1613 0.03087 0.02394 -0.0380 0.0465 0.8254
11.250 1.1720 0.03200 0.02517 -0.0369 0.0454 0.8349
11.500 1.1817 0.03321 0.02650 -0.0357 0.0444 0.8498
11.750 1.2010 0.03446 0.02795 -0.0366 0.0431 0.8947
12.000 1.2171 0.03630 0.02988 -0.0379 0.0418 1.0000
12.250 1.2185 0.03849 0.03210 -0.0366 0.0407 1.0000
12.500 1.2299 0.03987 0.03354 -0.0360 0.0400 1.0000
12.750 1.2424 0.04117 0.03491 -0.0355 0.0391 1.0000
13.000 1.2534 0.04262 0.03642 -0.0350 0.0382 1.0000
13.250 1.2637 0.04414 0.03800 -0.0345 0.0373 1.0000
13.500 1.2742 0.04568 0.03958 -0.0341 0.0362 1.0000
13.750 1.2838 0.04732 0.04125 -0.0337 0.0352 1.0000
14.000 1.2889 0.04939 0.04334 -0.0331 0.0340 1.0000
14.250 1.2946 0.05144 0.04545 -0.0326 0.0329 1.0000
14.500 1.3070 0.05295 0.04705 -0.0326 0.0317 1.0000
14.750 1.3180 0.05460 0.04877 -0.0326 0.0303 1.0000
15.000 1.3282 0.05635 0.05055 -0.0326 0.0289 1.0000
15.250 1.3313 0.05886 0.05307 -0.0324 0.0275 1.0000
15.500 1.3426 0.06055 0.05487 -0.0325 0.0257 1.0000
15.750 1.3539 0.06229 0.05665 -0.0327 0.0236 1.0000
16.000 1.3597 0.06465 0.05904 -0.0328 0.0215 1.0000
16.250 1.3680 0.06673 0.06119 -0.0329 0.0195 1.0000
16.500 1.3709 0.06950 0.06396 -0.0332 0.0181 1.0000
16.750 1.3746 0.07217 0.06673 -0.0333 0.0169 1.0000
17.000 1.3786 0.07488 0.06950 -0.0337 0.0158 1.0000
17.250 1.3798 0.07799 0.07266 -0.0341 0.0150 1.0000
17.500 1.3779 0.08157 0.07629 -0.0347 0.0144 1.0000
17.750 1.3791 0.08475 0.07958 -0.0353 0.0139 1.0000
18.000 1.3782 0.08828 0.08321 -0.0360 0.0134 1.0000
18.250 1.3788 0.09165 0.08666 -0.0369 0.0128 1.0000
18.500 1.3770 0.09542 0.09052 -0.0379 0.0125 1.0000
18.750 1.3744 0.09939 0.09455 -0.0391 0.0121 1.0000
19.000 1.3685 0.10387 0.09911 -0.0405 0.0118 1.0000
19.250 1.3629 0.10836 0.10369 -0.0421 0.0116 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 60-160 AIRFOIL (fx60160-il)