FX 60-157 AIRFOIL (fx60157-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 60-157 AIRFOIL (fx60157-il) Reynolds number: 100,000 Max Cl/Cd: 46.49 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx60157-il-100000.txt Download as CSV file: xf-fx60157-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-157 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5073 0.09979 0.09548 -0.0340 1.0000 0.1596
-9.000 -0.5217 0.09649 0.09224 -0.0335 1.0000 0.1648
-7.750 -0.7785 0.05522 0.04879 -0.0326 1.0000 0.0626
-7.500 -0.7690 0.05114 0.04459 -0.0314 1.0000 0.0608
-7.250 -0.7595 0.04710 0.04014 -0.0303 1.0000 0.0588
-7.000 -0.7459 0.04328 0.03579 -0.0292 1.0000 0.0575
-6.750 -0.7290 0.04033 0.03239 -0.0282 1.0000 0.0577
-6.500 -0.7101 0.03796 0.02965 -0.0273 1.0000 0.0591
-6.250 -0.6894 0.03645 0.02754 -0.0264 1.0000 0.0633
-6.000 -0.6709 0.03431 0.02552 -0.0257 1.0000 0.0682
-5.750 -0.6494 0.03327 0.02410 -0.0249 1.0000 0.0739
-5.500 -0.6306 0.03189 0.02286 -0.0243 1.0000 0.0813
-5.250 -0.6095 0.03088 0.02164 -0.0233 1.0000 0.0868
-5.000 -0.5899 0.02975 0.02062 -0.0221 1.0000 0.0903
-4.750 -0.5697 0.02898 0.01981 -0.0209 1.0000 0.0946
-4.500 -0.5498 0.02819 0.01903 -0.0199 1.0000 0.1006
-4.250 -0.5298 0.02747 0.01844 -0.0190 1.0000 0.1079
-4.000 -0.5085 0.02668 0.01777 -0.0186 1.0000 0.1182
-3.750 -0.4846 0.02592 0.01714 -0.0188 1.0000 0.1401
-3.500 -0.4578 0.02329 0.01749 -0.0203 1.0000 0.5522
-3.250 -0.4300 0.02594 0.02021 -0.0177 0.9908 0.6671
-3.000 -0.4007 0.02838 0.02250 -0.0156 0.9781 0.7021
-2.750 -0.3736 0.03038 0.02436 -0.0135 0.9685 0.7283
-2.500 -0.3592 0.03155 0.02547 -0.0091 0.9586 0.7464
-2.250 -0.3422 0.03257 0.02640 -0.0055 0.9489 0.7634
-2.000 -0.3212 0.03385 0.02759 -0.0020 0.9410 0.7829
-1.750 -0.3141 0.03429 0.02800 0.0034 0.9309 0.8025
-1.500 -0.2967 0.03503 0.02865 0.0072 0.9220 0.8255
-1.250 -0.2786 0.03532 0.02887 0.0107 0.9119 0.8438
-1.000 -0.2577 0.03525 0.02869 0.0119 0.9012 0.8572
-0.750 -0.2183 0.03541 0.02871 0.0099 0.8942 0.8652
-0.500 -0.2026 0.03506 0.02828 0.0111 0.8834 0.8735
-0.250 -0.1592 0.03522 0.02830 0.0082 0.8776 0.8801
0.000 -0.1451 0.03485 0.02786 0.0094 0.8660 0.8875
0.250 -0.1125 0.03479 0.02772 0.0084 0.8583 0.8934
0.500 -0.0845 0.03459 0.02744 0.0076 0.8486 0.9001
0.750 -0.0580 0.03435 0.02713 0.0073 0.8378 0.9054
1.000 -0.0134 0.03417 0.02687 0.0043 0.8309 0.9105
1.250 0.0020 0.03395 0.02662 0.0056 0.8192 0.9175
1.500 0.0517 0.03376 0.02637 0.0018 0.8143 0.9219
1.750 0.0696 0.03356 0.02615 0.0027 0.8019 0.9272
2.000 0.0968 0.03327 0.02583 0.0024 0.7918 0.9322
2.250 0.1432 0.03282 0.02536 -0.0008 0.7842 0.9356
2.500 0.1696 0.03259 0.02512 -0.0011 0.7726 0.9396
2.750 0.2162 0.03200 0.02452 -0.0041 0.7671 0.9429
3.000 0.2388 0.03180 0.02433 -0.0038 0.7548 0.9473
3.250 0.2966 0.03092 0.02347 -0.0083 0.7505 0.9498
3.500 0.3230 0.03061 0.02319 -0.0085 0.7377 0.9541
3.750 0.3777 0.02946 0.02207 -0.0122 0.7342 0.9565
4.000 0.4008 0.02908 0.02172 -0.0117 0.7211 0.9599
4.250 0.4622 0.02751 0.02022 -0.0160 0.7182 0.9612
4.500 0.4949 0.02688 0.01964 -0.0168 0.7055 0.9641
4.750 0.5616 0.02517 0.01802 -0.0220 0.7022 0.9656
5.000 0.5961 0.02444 0.01734 -0.0229 0.6892 0.9690
5.250 0.6436 0.02338 0.01636 -0.0256 0.6782 0.9714
5.500 0.7143 0.02194 0.01497 -0.0320 0.6659 0.9719
5.750 0.7586 0.02131 0.01439 -0.0347 0.6453 0.9738
6.000 0.8129 0.02067 0.01374 -0.0390 0.6221 0.9752
6.250 0.8537 0.02045 0.01349 -0.0414 0.5957 0.9777
6.500 0.8937 0.02038 0.01332 -0.0437 0.5670 0.9804
6.750 0.9266 0.02049 0.01332 -0.0448 0.5375 0.9836
7.000 0.9530 0.02079 0.01357 -0.0451 0.5073 0.9872
7.250 0.9771 0.02115 0.01383 -0.0450 0.4801 0.9912
7.500 1.0038 0.02159 0.01413 -0.0455 0.4534 0.9954
7.750 1.0124 0.02196 0.01450 -0.0427 0.4313 1.0000
8.000 1.0095 0.02218 0.01467 -0.0377 0.4149 1.0000
8.250 1.0126 0.02253 0.01496 -0.0340 0.3971 1.0000
8.500 1.0230 0.02302 0.01542 -0.0318 0.3777 1.0000
8.750 1.0376 0.02360 0.01593 -0.0304 0.3579 1.0000
9.000 1.0560 0.02430 0.01652 -0.0298 0.3380 1.0000
9.250 1.0706 0.02506 0.01729 -0.0288 0.3186 1.0000
9.500 1.0871 0.02587 0.01811 -0.0282 0.3006 1.0000
9.750 1.1046 0.02674 0.01898 -0.0278 0.2833 1.0000
10.000 1.1218 0.02767 0.01992 -0.0274 0.2671 1.0000
10.250 1.1393 0.02864 0.02091 -0.0271 0.2519 1.0000
10.500 1.1565 0.02965 0.02192 -0.0269 0.2377 1.0000
10.750 1.1743 0.03073 0.02296 -0.0268 0.2239 1.0000
11.000 1.1864 0.03189 0.02429 -0.0260 0.2105 1.0000
11.250 1.1994 0.03312 0.02563 -0.0255 0.1978 1.0000
11.500 1.2131 0.03443 0.02700 -0.0251 0.1861 1.0000
11.750 1.2261 0.03584 0.02842 -0.0247 0.1745 1.0000
12.000 1.2355 0.03738 0.02999 -0.0241 0.1623 1.0000
12.250 1.2400 0.03914 0.03194 -0.0232 0.1500 1.0000
12.500 1.2443 0.04109 0.03398 -0.0224 0.1378 1.0000
12.750 1.2478 0.04333 0.03626 -0.0217 0.1252 1.0000
13.000 1.2493 0.04586 0.03883 -0.0210 0.1128 1.0000
13.250 1.2505 0.04853 0.04150 -0.0204 0.1020 1.0000
13.500 1.2552 0.05113 0.04397 -0.0201 0.0927 1.0000
13.750 1.2598 0.05400 0.04704 -0.0197 0.0852 1.0000
14.000 1.2732 0.05643 0.04936 -0.0197 0.0791 1.0000
14.250 1.2781 0.05938 0.05261 -0.0196 0.0751 1.0000
14.500 1.2875 0.06192 0.05523 -0.0197 0.0716 1.0000
14.750 1.3069 0.06411 0.05733 -0.0200 0.0685 1.0000
15.000 1.3084 0.06754 0.06108 -0.0201 0.0669 1.0000
15.250 1.3096 0.07113 0.06495 -0.0203 0.0654 1.0000
15.500 1.3092 0.07497 0.06906 -0.0207 0.0643 1.0000
15.750 1.3058 0.07916 0.07349 -0.0213 0.0635 1.0000
16.000 1.2987 0.08381 0.07840 -0.0223 0.0629 1.0000
16.250 1.2868 0.08910 0.08396 -0.0238 0.0625 1.0000
16.500 1.2692 0.09532 0.09046 -0.0260 0.0623 1.0000
16.750 1.2429 0.10304 0.09850 -0.0296 0.0625 1.0000
17.000 1.1973 0.11445 0.11028 -0.0361 0.0633 1.0000
17.250 1.1033 0.13742 0.13366 -0.0516 0.0660 1.0000
17.500 1.0239 0.16306 0.15934 -0.0679 0.0687 1.0000
17.750 0.9521 0.19586 0.19194 -0.0858 0.0792 1.0000
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