FX 66-S-171 AIRFOIL (fx66s171-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 66-S-171 AIRFOIL (fx66s171-il) Reynolds number: 500,000 Max Cl/Cd: 113.38 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx66s171-il-500000-n5.txt Download as CSV file: xf-fx66s171-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-S-171 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.250 -0.6591 0.08374 0.08081 -0.0572 1.0000 0.0071
-15.000 -0.6837 0.07524 0.07213 -0.0622 1.0000 0.0071
-14.750 -0.7057 0.06776 0.06450 -0.0667 1.0000 0.0071
-14.500 -0.7193 0.06217 0.05876 -0.0699 1.0000 0.0071
-14.250 -0.7297 0.05746 0.05394 -0.0726 1.0000 0.0071
-14.000 -0.7398 0.05307 0.04944 -0.0749 1.0000 0.0072
-13.750 -0.7026 0.04653 0.04260 -0.0895 0.9246 0.0073
-13.500 -0.6528 0.04127 0.03675 -0.1049 0.8320 0.0075
-13.250 -0.6501 0.03940 0.03460 -0.1053 0.7856 0.0077
-13.000 -0.6531 0.03695 0.03191 -0.1054 0.7580 0.0077
-12.750 -0.6512 0.03495 0.02973 -0.1056 0.7381 0.0079
-12.500 -0.6473 0.03316 0.02777 -0.1057 0.7217 0.0081
-12.250 -0.6458 0.03116 0.02560 -0.1055 0.7081 0.0082
-12.000 -0.6424 0.02938 0.02366 -0.1053 0.6960 0.0083
-11.750 -0.6372 0.02779 0.02193 -0.1050 0.6850 0.0085
-11.500 -0.6327 0.02625 0.02024 -0.1044 0.6754 0.0086
-11.250 -0.6288 0.02483 0.01868 -0.1035 0.6662 0.0087
-11.000 -0.6255 0.02368 0.01740 -0.1018 0.6582 0.0088
-10.750 -0.6204 0.02280 0.01639 -0.0996 0.6505 0.0088
-10.500 -0.6119 0.02170 0.01520 -0.0980 0.6438 0.0091
-10.250 -0.5982 0.02088 0.01427 -0.0969 0.6367 0.0093
-10.000 -0.5815 0.02019 0.01349 -0.0959 0.6303 0.0096
-9.750 -0.5633 0.01954 0.01276 -0.0951 0.6238 0.0099
-9.500 -0.5442 0.01894 0.01206 -0.0943 0.6179 0.0102
-9.250 -0.5240 0.01834 0.01137 -0.0936 0.6126 0.0106
-9.000 -0.5031 0.01777 0.01071 -0.0930 0.6071 0.0109
-8.500 -0.4595 0.01670 0.00944 -0.0919 0.5975 0.0119
-8.250 -0.4372 0.01616 0.00885 -0.0915 0.5927 0.0125
-8.000 -0.4135 0.01577 0.00838 -0.0911 0.5882 0.0134
-7.750 -0.3892 0.01538 0.00793 -0.0908 0.5842 0.0144
-7.500 -0.3646 0.01500 0.00749 -0.0905 0.5799 0.0154
-7.250 -0.3402 0.01459 0.00702 -0.0903 0.5758 0.0170
-7.000 -0.3151 0.01428 0.00663 -0.0900 0.5721 0.0184
-6.750 -0.2895 0.01394 0.00623 -0.0898 0.5686 0.0200
-6.500 -0.2639 0.01360 0.00587 -0.0897 0.5648 0.0221
-6.250 -0.2378 0.01333 0.00553 -0.0896 0.5610 0.0242
-6.000 -0.2118 0.01305 0.00521 -0.0895 0.5574 0.0273
-5.750 -0.1852 0.01281 0.00494 -0.0894 0.5542 0.0311
-5.500 -0.1587 0.01252 0.00465 -0.0894 0.5511 0.0375
-5.250 -0.1321 0.01225 0.00438 -0.0894 0.5480 0.0458
-5.000 -0.1054 0.01199 0.00412 -0.0893 0.5450 0.0583
-4.750 -0.0788 0.01173 0.00388 -0.0893 0.5424 0.0751
-4.500 -0.0523 0.01140 0.00364 -0.0893 0.5399 0.1029
-4.250 -0.0259 0.01101 0.00338 -0.0894 0.5372 0.1450
-4.000 0.0002 0.01057 0.00312 -0.0895 0.5342 0.1999
-3.750 0.0255 0.01001 0.00286 -0.0895 0.5313 0.2888
-3.500 0.0523 0.00974 0.00272 -0.0896 0.5286 0.3422
-3.000 0.1077 0.00952 0.00264 -0.0897 0.5240 0.4099
-2.750 0.1363 0.00951 0.00262 -0.0899 0.5217 0.4291
-2.500 0.1646 0.00951 0.00262 -0.0900 0.5193 0.4481
-2.250 0.1931 0.00954 0.00263 -0.0901 0.5169 0.4620
-2.000 0.2214 0.00957 0.00262 -0.0902 0.5146 0.4718
-1.750 0.2498 0.00962 0.00262 -0.0904 0.5127 0.4800
-1.500 0.2782 0.00966 0.00262 -0.0905 0.5108 0.4864
-1.250 0.3070 0.00969 0.00260 -0.0907 0.5091 0.4905
-1.000 0.3356 0.00969 0.00260 -0.0909 0.5072 0.4937
-0.750 0.3642 0.00970 0.00260 -0.0911 0.5053 0.4972
-0.500 0.3927 0.00973 0.00261 -0.0913 0.5033 0.5008
-0.250 0.4211 0.00977 0.00261 -0.0914 0.5012 0.5044
0.000 0.4494 0.00982 0.00262 -0.0916 0.4989 0.5075
0.250 0.4774 0.00987 0.00264 -0.0917 0.4965 0.5106
0.500 0.5057 0.00989 0.00268 -0.0918 0.4944 0.5142
0.750 0.5342 0.00992 0.00272 -0.0920 0.4923 0.5180
1.000 0.5627 0.00997 0.00275 -0.0922 0.4901 0.5217
1.250 0.5908 0.00999 0.00280 -0.0924 0.4880 0.5249
1.500 0.6190 0.01004 0.00286 -0.0925 0.4860 0.5283
1.750 0.6470 0.01010 0.00291 -0.0926 0.4840 0.5320
2.000 0.6749 0.01018 0.00297 -0.0928 0.4820 0.5359
2.250 0.7029 0.01024 0.00305 -0.0929 0.4801 0.5395
2.500 0.7310 0.01027 0.00313 -0.0930 0.4781 0.5437
2.750 0.7590 0.01032 0.00322 -0.0932 0.4760 0.5484
3.000 0.7870 0.01038 0.00331 -0.0933 0.4739 0.5533
3.250 0.8147 0.01044 0.00341 -0.0934 0.4719 0.5583
3.500 0.8424 0.01050 0.00350 -0.0935 0.4699 0.5637
3.750 0.8699 0.01059 0.00361 -0.0936 0.4680 0.5689
4.000 0.8972 0.01068 0.00372 -0.0937 0.4661 0.5750
4.250 0.9248 0.01074 0.00386 -0.0938 0.4641 0.5819
4.500 0.9523 0.01079 0.00400 -0.0939 0.4618 0.5884
4.750 0.9796 0.01086 0.00413 -0.0939 0.4592 0.5953
5.000 1.0067 0.01094 0.00426 -0.0939 0.4566 0.6022
5.250 1.0335 0.01102 0.00440 -0.0939 0.4542 0.6104
5.500 1.0600 0.01113 0.00454 -0.0939 0.4517 0.6190
5.750 1.0869 0.01120 0.00471 -0.0939 0.4488 0.6290
6.000 1.1136 0.01127 0.00489 -0.0938 0.4459 0.6397
6.250 1.1399 0.01135 0.00507 -0.0937 0.4430 0.6525
6.500 1.1659 0.01144 0.00525 -0.0936 0.4404 0.6678
6.750 1.1915 0.01155 0.00543 -0.0934 0.4378 0.6855
7.000 1.2171 0.01163 0.00565 -0.0932 0.4350 0.7062
7.250 1.2425 0.01169 0.00588 -0.0929 0.4315 0.7332
7.500 1.2667 0.01173 0.00610 -0.0924 0.4281 0.7710
8.000 1.3183 0.01166 0.00652 -0.0919 0.4191 1.0000
8.250 1.3402 0.01182 0.00667 -0.0911 0.4066 1.0000
8.500 1.3586 0.01210 0.00689 -0.0896 0.3841 1.0000
8.750 1.3719 0.01263 0.00729 -0.0874 0.3559 1.0000
9.000 1.3808 0.01335 0.00785 -0.0846 0.3261 1.0000
9.250 1.3857 0.01410 0.00849 -0.0810 0.3022 1.0000
9.500 1.3869 0.01496 0.00924 -0.0770 0.2751 1.0000
9.750 1.3800 0.01621 0.01033 -0.0721 0.2431 1.0000
10.000 1.3673 0.01787 0.01183 -0.0670 0.2104 1.0000
10.250 1.3537 0.01991 0.01375 -0.0627 0.1829 1.0000
10.500 1.3391 0.02243 0.01618 -0.0593 0.1572 1.0000
10.750 1.3251 0.02533 0.01900 -0.0567 0.1342 1.0000
11.000 1.3111 0.02855 0.02214 -0.0547 0.1128 1.0000
11.250 1.2961 0.03206 0.02557 -0.0530 0.0908 1.0000
11.500 1.2872 0.03517 0.02865 -0.0517 0.0765 1.0000
11.750 1.2771 0.03845 0.03189 -0.0506 0.0620 1.0000
12.000 1.2672 0.04179 0.03521 -0.0496 0.0483 1.0000
12.250 1.2617 0.04482 0.03823 -0.0489 0.0395 1.0000
12.500 1.2573 0.04786 0.04127 -0.0484 0.0315 1.0000
12.750 1.2547 0.05081 0.04422 -0.0480 0.0255 1.0000
13.000 1.2535 0.05369 0.04711 -0.0477 0.0213 1.0000
13.250 1.2550 0.05636 0.04983 -0.0476 0.0186 1.0000
13.500 1.2555 0.05918 0.05269 -0.0475 0.0164 1.0000
13.750 1.2585 0.06178 0.05534 -0.0475 0.0149 1.0000
14.000 1.2597 0.06465 0.05825 -0.0476 0.0133 1.0000
14.250 1.2623 0.06736 0.06101 -0.0477 0.0125 1.0000
14.500 1.2650 0.07013 0.06386 -0.0478 0.0115 1.0000
14.750 1.2674 0.07295 0.06673 -0.0481 0.0106 1.0000
15.000 1.2686 0.07600 0.06983 -0.0484 0.0099 1.0000
15.250 1.2711 0.07891 0.07281 -0.0488 0.0093 1.0000
15.500 1.2741 0.08179 0.07576 -0.0492 0.0087 1.0000
15.750 1.2768 0.08474 0.07877 -0.0496 0.0083 1.0000
16.000 1.2785 0.08787 0.08196 -0.0502 0.0077 1.0000
16.250 1.2795 0.09113 0.08528 -0.0508 0.0074 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 66-S-171 AIRFOIL (fx66s171-il)