FX 66-S-171 AIRFOIL (fx66s171-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 66-S-171 AIRFOIL (fx66s171-il) Reynolds number: 50,000 Max Cl/Cd: 4.98 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx66s171-il-50000.txt Download as CSV file: xf-fx66s171-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.2457 0.11891 0.11329 -0.0381 1.0000 0.2950
-10.000 -0.2483 0.11765 0.11227 -0.0350 1.0000 0.2990
-9.750 -0.3114 0.12460 0.11956 -0.0245 1.0000 0.2968
-9.500 -0.3118 0.12242 0.11739 -0.0307 0.9896 0.3087
-9.250 -0.2636 0.11457 0.10944 -0.0359 0.9818 0.3145
-9.000 -0.3451 0.09835 0.09335 -0.0578 0.9707 0.1725
-8.750 -0.3729 0.08598 0.08101 -0.0691 0.9599 0.1411
-8.500 -0.4316 0.07562 0.07054 -0.0803 0.9452 0.1266
-8.250 -0.4398 0.07025 0.06502 -0.0840 0.9339 0.1234
-8.000 -0.4629 0.06407 0.05842 -0.0880 0.9235 0.1184
-7.750 -0.4820 0.05984 0.05332 -0.0892 0.9132 0.1142
-7.500 -0.4782 0.05678 0.04982 -0.0890 0.9048 0.1144
-7.250 -0.4567 0.05312 0.04604 -0.0904 0.8981 0.1173
-7.000 -0.4431 0.05057 0.04317 -0.0903 0.8910 0.1189
-6.750 -0.4271 0.04809 0.04029 -0.0901 0.8840 0.1209
-6.500 -0.3977 0.04557 0.03718 -0.0917 0.8788 0.1268
-6.250 -0.3954 0.04453 0.03586 -0.0889 0.8724 0.1303
-6.000 -0.3679 0.04276 0.03403 -0.0893 0.8673 0.1380
-5.750 -0.3416 0.04129 0.03244 -0.0894 0.8624 0.1500
-5.500 -0.3331 0.04064 0.03160 -0.0868 0.8567 0.1598
-5.250 -0.3097 0.03946 0.03062 -0.0863 0.8522 0.1792
-5.000 -0.2825 0.03800 0.02939 -0.0866 0.8484 0.2175
-4.750 -0.2851 0.03744 0.02931 -0.0830 0.8447 0.2548
-4.500 -0.2844 0.03754 0.03087 -0.0782 0.8409 0.4072
-4.250 -0.2789 0.04074 0.03415 -0.0711 0.8371 0.5100
-4.000 -0.2641 0.04319 0.03653 -0.0654 0.8337 0.5580
-3.750 -0.2660 0.04449 0.03777 -0.0599 0.8315 0.5858
-3.500 -0.2671 0.04567 0.03891 -0.0542 0.8295 0.6109
-3.250 -0.2664 0.04665 0.03984 -0.0488 0.8282 0.6357
-3.000 -0.2642 0.04735 0.04044 -0.0445 0.8277 0.6619
-2.750 -0.2640 0.04803 0.04109 -0.0391 0.8277 0.6842
-2.500 -0.4348 0.04130 0.03484 -0.0247 1.0000 0.6010
-2.250 -0.4284 0.04223 0.03571 -0.0199 1.0000 0.6333
-2.000 -0.4218 0.04292 0.03634 -0.0155 1.0000 0.6638
-1.750 -0.4155 0.04344 0.03679 -0.0110 1.0000 0.6938
-1.500 -0.4094 0.04381 0.03709 -0.0067 1.0000 0.7234
-1.250 -0.4051 0.04397 0.03721 -0.0020 1.0000 0.7493
-1.000 -0.3968 0.04405 0.03718 0.0013 1.0000 0.7738
-0.750 -0.3858 0.04413 0.03712 0.0035 1.0000 0.7947
-0.500 -0.3706 0.04429 0.03710 0.0043 1.0000 0.8124
-0.250 -0.3536 0.04450 0.03714 0.0045 1.0000 0.8273
0.000 -0.3354 0.04479 0.03727 0.0042 1.0000 0.8415
0.250 -0.3083 0.04570 0.03799 0.0021 0.9961 0.8559
0.500 -0.2751 0.04701 0.03912 -0.0013 0.9877 0.8716
0.750 -0.2416 0.04840 0.04037 -0.0048 0.9785 0.8894
1.000 -0.2022 0.05018 0.04202 -0.0095 0.9684 0.9097
1.250 -0.1510 0.05247 0.04421 -0.0168 0.9563 0.9341
1.500 -0.1000 0.05424 0.04588 -0.0248 0.9423 0.9646
1.750 -0.0630 0.05557 0.04704 -0.0303 0.9277 1.0000
2.000 -0.0331 0.05678 0.04806 -0.0343 0.9141 1.0000
2.250 -0.0026 0.05826 0.04936 -0.0383 0.9017 1.0000
2.500 0.0346 0.06065 0.05153 -0.0435 0.8932 1.0000
2.750 0.0702 0.06265 0.05335 -0.0480 0.8809 1.0000
3.000 0.0955 0.06398 0.05451 -0.0505 0.8678 1.0000
3.250 0.1209 0.06563 0.05600 -0.0530 0.8564 1.0000
3.500 0.1630 0.06901 0.05918 -0.0581 0.8489 1.0000
3.750 0.1800 0.06986 0.05991 -0.0588 0.8358 1.0000
4.000 0.1985 0.07126 0.06118 -0.0596 0.8242 1.0000
4.250 0.2292 0.07400 0.06378 -0.0624 0.8168 1.0000
4.500 0.2500 0.07551 0.06519 -0.0633 0.8051 1.0000
4.750 0.2639 0.07684 0.06644 -0.0634 0.7941 1.0000
5.000 0.3024 0.08051 0.07000 -0.0669 0.7876 1.0000
5.250 0.3086 0.08094 0.07039 -0.0657 0.7753 1.0000
5.500 0.3225 0.08261 0.07200 -0.0658 0.7666 1.0000
5.750 0.3520 0.08535 0.07468 -0.0678 0.7581 1.0000
6.000 0.3594 0.08642 0.07573 -0.0670 0.7473 1.0000
6.250 0.3945 0.09008 0.07936 -0.0697 0.7415 1.0000
6.500 0.3956 0.09051 0.07978 -0.0682 0.7303 1.0000
6.750 0.4197 0.09347 0.08272 -0.0696 0.7243 1.0000
7.000 0.4321 0.09485 0.08409 -0.0694 0.7132 1.0000
7.250 0.4430 0.09676 0.08603 -0.0693 0.7056 1.0000
7.500 0.4662 0.09927 0.08855 -0.0704 0.6969 1.0000
7.750 0.4723 0.10083 0.09013 -0.0698 0.6881 1.0000
8.000 0.4994 0.10388 0.09320 -0.0714 0.6807 1.0000
8.250 0.5016 0.10515 0.09450 -0.0705 0.6714 1.0000
8.500 0.5317 0.10870 0.09811 -0.0723 0.6647 1.0000
8.750 0.5303 0.10965 0.09909 -0.0712 0.6549 1.0000
9.000 0.5617 0.11359 0.10310 -0.0732 0.6489 1.0000
9.250 0.5581 0.11431 0.10387 -0.0720 0.6385 1.0000
9.500 0.5918 0.11879 0.10844 -0.0742 0.6331 1.0000
9.750 0.5848 0.11914 0.10884 -0.0729 0.6224 1.0000
10.000 0.6157 0.12367 0.11345 -0.0748 0.6172 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 66-S-171 AIRFOIL (fx66s171-il)