NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63-210 AIRFOIL (n63210-il) Reynolds number: 50,000 Max Cl/Cd: 32 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63210-il-50000-n5.txt Download as CSV file: xf-n63210-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-210 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5648 0.09670 0.08964 -0.0205 1.0000 0.0549
-9.750 -0.5695 0.09094 0.08395 -0.0241 1.0000 0.0546
-9.500 -0.5765 0.08452 0.07758 -0.0288 1.0000 0.0539
-9.250 -0.5893 0.07823 0.07133 -0.0336 1.0000 0.0529
-9.000 -0.6067 0.07286 0.06596 -0.0371 1.0000 0.0522
-8.750 -0.6218 0.06813 0.06114 -0.0390 1.0000 0.0517
-8.500 -0.6300 0.06383 0.05669 -0.0402 1.0000 0.0522
-8.250 -0.6341 0.05976 0.05239 -0.0407 1.0000 0.0532
-8.000 -0.6342 0.05575 0.04808 -0.0409 1.0000 0.0541
-7.750 -0.6305 0.05177 0.04373 -0.0407 1.0000 0.0548
-7.500 -0.6229 0.04803 0.03957 -0.0402 1.0000 0.0554
-7.250 -0.6124 0.04479 0.03579 -0.0395 1.0000 0.0577
-7.000 -0.5985 0.04185 0.03226 -0.0386 1.0000 0.0591
-6.750 -0.5813 0.03903 0.02899 -0.0376 1.0000 0.0594
-6.500 -0.5622 0.03656 0.02605 -0.0366 1.0000 0.0598
-6.250 -0.5421 0.03410 0.02339 -0.0356 1.0000 0.0605
-6.000 -0.5212 0.03208 0.02119 -0.0346 1.0000 0.0614
-5.750 -0.5000 0.03039 0.01936 -0.0334 1.0000 0.0626
-5.500 -0.4789 0.02895 0.01781 -0.0321 1.0000 0.0640
-5.250 -0.4583 0.02770 0.01646 -0.0306 1.0000 0.0658
-5.000 -0.4387 0.02662 0.01522 -0.0290 1.0000 0.0679
-4.750 -0.4201 0.02570 0.01415 -0.0273 1.0000 0.0716
-4.500 -0.4030 0.02475 0.01317 -0.0258 1.0000 0.0763
-4.250 -0.3861 0.02388 0.01225 -0.0243 1.0000 0.0830
-4.000 -0.3690 0.02300 0.01131 -0.0230 1.0000 0.0901
-3.750 -0.3512 0.02207 0.01037 -0.0218 1.0000 0.1024
-3.500 -0.3334 0.02075 0.00931 -0.0211 1.0000 0.1420
-3.250 -0.3207 0.01853 0.00876 -0.0201 1.0000 0.4172
-3.000 -0.3123 0.01831 0.00930 -0.0157 1.0000 0.6166
-2.750 -0.3128 0.01879 0.01026 -0.0079 1.0000 0.7215
-2.500 -0.3042 0.01940 0.01096 -0.0017 0.9941 0.7901
-2.250 -0.2746 0.01961 0.01098 -0.0012 0.9843 0.8212
-2.000 -0.2395 0.01965 0.01074 -0.0025 0.9751 0.8399
-1.750 -0.2043 0.01965 0.01052 -0.0042 0.9656 0.8571
-1.500 -0.1678 0.01964 0.01031 -0.0061 0.9565 0.8738
-1.250 -0.1270 0.01966 0.01013 -0.0090 0.9486 0.8898
-1.000 -0.0887 0.01963 0.00996 -0.0114 0.9392 0.9063
-0.750 -0.0457 0.01962 0.00982 -0.0148 0.9311 0.9220
-0.500 0.0018 0.01962 0.00969 -0.0191 0.9235 0.9368
-0.250 0.0505 0.01960 0.00959 -0.0238 0.9156 0.9510
0.000 0.1037 0.01957 0.00948 -0.0293 0.9087 0.9638
0.250 0.1545 0.01952 0.00940 -0.0345 0.9003 0.9767
0.500 0.2092 0.01943 0.00931 -0.0403 0.8934 0.9885
0.750 0.2489 0.01942 0.00931 -0.0435 0.8829 1.0000
1.000 0.2732 0.01949 0.00938 -0.0436 0.8709 1.0000
1.250 0.2972 0.01961 0.00952 -0.0435 0.8592 1.0000
1.500 0.3218 0.01976 0.00969 -0.0435 0.8482 1.0000
1.750 0.3486 0.01991 0.00988 -0.0436 0.8378 1.0000
2.000 0.3705 0.02018 0.01018 -0.0430 0.8253 1.0000
2.250 0.3939 0.02046 0.01055 -0.0426 0.8132 1.0000
2.500 0.4186 0.02075 0.01091 -0.0424 0.8013 1.0000
2.750 0.4450 0.02101 0.01127 -0.0422 0.7899 1.0000
3.000 0.4734 0.02121 0.01162 -0.0423 0.7790 1.0000
3.250 0.4998 0.02146 0.01201 -0.0419 0.7666 1.0000
3.500 0.5252 0.02170 0.01242 -0.0414 0.7530 1.0000
3.750 0.5514 0.02188 0.01278 -0.0407 0.7385 1.0000
4.000 0.5783 0.02195 0.01308 -0.0399 0.7227 1.0000
4.250 0.6022 0.02197 0.01331 -0.0384 0.7015 1.0000
4.500 0.6249 0.02103 0.01244 -0.0345 0.6561 1.0000
4.750 0.6383 0.01995 0.01094 -0.0284 0.5493 1.0000
5.000 0.6491 0.02045 0.01044 -0.0241 0.3581 1.0000
5.250 0.6545 0.02288 0.01150 -0.0217 0.1610 1.0000
5.500 0.6676 0.02492 0.01302 -0.0205 0.1009 1.0000
5.750 0.6845 0.02644 0.01448 -0.0193 0.0842 1.0000
6.250 0.7212 0.02916 0.01735 -0.0171 0.0716 1.0000
6.500 0.7418 0.03049 0.01881 -0.0161 0.0679 1.0000
6.750 0.7646 0.03191 0.02035 -0.0153 0.0651 1.0000
7.000 0.7913 0.03364 0.02207 -0.0148 0.0629 1.0000
7.250 0.8208 0.03533 0.02402 -0.0145 0.0612 1.0000
7.500 0.8480 0.03719 0.02618 -0.0142 0.0584 1.0000
7.750 0.8723 0.03916 0.02842 -0.0138 0.0554 1.0000
8.000 0.8956 0.04147 0.03099 -0.0133 0.0536 1.0000
8.250 0.9170 0.04412 0.03403 -0.0128 0.0527 1.0000
8.500 0.9356 0.04706 0.03732 -0.0121 0.0520 1.0000
8.750 0.9508 0.05026 0.04089 -0.0113 0.0515 1.0000
9.000 0.9624 0.05374 0.04475 -0.0103 0.0511 1.0000
9.250 0.9699 0.05745 0.04886 -0.0092 0.0509 1.0000
9.500 0.9731 0.06133 0.05313 -0.0081 0.0508 1.0000
9.750 0.9715 0.06534 0.05751 -0.0070 0.0508 1.0000
10.000 0.9667 0.06940 0.06185 -0.0060 0.0505 1.0000
10.500 0.9429 0.07769 0.07050 -0.0045 0.0498 1.0000
10.750 0.9252 0.08190 0.07489 -0.0046 0.0498 1.0000
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Polar data table (+)
Polar graphs
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