NACA 0010-65 (naca001065-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 0010-65 (naca001065-il) Reynolds number: 500,000 Max Cl/Cd: 43.08 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca001065-il-500000.txt Download as CSV file: xf-naca001065-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 0010-65
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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.7770 0.06475 0.06220 -0.0531 1.0000 0.0193
-11.500 -0.8229 0.05753 0.05480 -0.0525 1.0000 0.0188
-11.250 -0.8745 0.05153 0.04856 -0.0469 1.0000 0.0185
-11.000 -0.9228 0.04749 0.04429 -0.0377 1.0000 0.0183
-10.750 -0.9612 0.04428 0.04084 -0.0283 1.0000 0.0183
-10.500 -0.9866 0.04099 0.03725 -0.0207 1.0000 0.0185
-10.250 -1.0045 0.03793 0.03387 -0.0139 1.0000 0.0188
-10.000 -1.0159 0.03504 0.03062 -0.0078 1.0000 0.0193
-9.750 -1.0197 0.03268 0.02792 -0.0027 1.0000 0.0198
-9.500 -1.0161 0.03119 0.02612 0.0014 1.0000 0.0203
-9.250 -1.0171 0.02831 0.02285 0.0060 1.0000 0.0209
-9.000 -1.0070 0.02693 0.02139 0.0087 1.0000 0.0216
-8.750 -0.9932 0.02646 0.02088 0.0110 1.0000 0.0222
-8.500 -0.9784 0.02604 0.02042 0.0131 1.0000 0.0230
-8.250 -0.9636 0.02537 0.01963 0.0153 1.0000 0.0240
-8.000 -0.9442 0.02445 0.01851 0.0166 0.9995 0.0251
-7.750 -0.9164 0.02387 0.01773 0.0163 0.9981 0.0260
-7.500 -0.8942 0.02126 0.01487 0.0168 0.9972 0.0275
-7.250 -0.8659 0.02069 0.01428 0.0161 0.9958 0.0288
-7.000 -0.8369 0.02020 0.01373 0.0154 0.9947 0.0304
-6.750 -0.8086 0.01955 0.01296 0.0150 0.9936 0.0318
-6.500 -0.7814 0.01900 0.01230 0.0149 0.9917 0.0328
-6.250 -0.7550 0.01781 0.01098 0.0149 0.9904 0.0342
-6.000 -0.7286 0.01690 0.01005 0.0148 0.9890 0.0360
-5.750 -0.6999 0.01639 0.00951 0.0143 0.9875 0.0374
-5.500 -0.6700 0.01593 0.00900 0.0136 0.9861 0.0390
-5.250 -0.6392 0.01553 0.00856 0.0126 0.9849 0.0406
-5.000 -0.6080 0.01515 0.00813 0.0116 0.9839 0.0417
-4.750 -0.5833 0.01456 0.00748 0.0121 0.9819 0.0430
-4.500 -0.5601 0.01398 0.00690 0.0128 0.9789 0.0454
-4.250 -0.5320 0.01366 0.00658 0.0125 0.9767 0.0478
-4.000 -0.5024 0.01337 0.00626 0.0119 0.9749 0.0502
-3.750 -0.4711 0.01312 0.00599 0.0110 0.9735 0.0527
-3.500 -0.4398 0.01275 0.00567 0.0100 0.9723 0.0605
-3.250 -0.4081 0.01241 0.00543 0.0089 0.9713 0.0811
-3.000 -0.3886 0.01203 0.00524 0.0104 0.9673 0.1193
-2.750 -0.3639 0.01162 0.00508 0.0107 0.9642 0.1742
-2.500 -0.3373 0.01110 0.00491 0.0105 0.9620 0.2594
-2.250 -0.3152 0.01002 0.00466 0.0110 0.9600 0.4611
-2.000 -0.3002 0.00866 0.00451 0.0135 0.9581 0.7421
-1.750 -0.2702 0.00852 0.00473 0.0135 0.9572 0.8398
-1.500 -0.2518 0.00864 0.00494 0.0160 0.9522 0.8778
-1.250 -0.2234 0.00882 0.00515 0.0162 0.9496 0.9025
-1.000 -0.1917 0.00908 0.00542 0.0158 0.9479 0.9185
-0.750 -0.1584 0.00936 0.00569 0.0150 0.9465 0.9303
-0.500 -0.1178 0.00976 0.00609 0.0126 0.9461 0.9378
-0.250 -0.0771 0.01012 0.00644 0.0101 0.9457 0.9449
0.000 -0.0002 0.01037 0.00669 0.0000 0.9462 0.9462
0.250 0.0773 0.01012 0.00644 -0.0102 0.9449 0.9457
0.500 0.1178 0.00976 0.00609 -0.0126 0.9377 0.9461
0.750 0.1583 0.00936 0.00569 -0.0150 0.9304 0.9465
1.000 0.1917 0.00908 0.00542 -0.0158 0.9183 0.9479
1.250 0.2234 0.00882 0.00515 -0.0162 0.9025 0.9496
1.500 0.2518 0.00863 0.00494 -0.0160 0.8778 0.9522
1.750 0.2702 0.00852 0.00473 -0.0135 0.8400 0.9572
2.000 0.3003 0.00865 0.00451 -0.0136 0.7432 0.9581
2.250 0.3152 0.01002 0.00466 -0.0110 0.4610 0.9601
2.500 0.3373 0.01110 0.00491 -0.0105 0.2596 0.9620
2.750 0.3638 0.01162 0.00508 -0.0107 0.1748 0.9642
3.000 0.3885 0.01203 0.00523 -0.0104 0.1195 0.9673
3.250 0.4080 0.01241 0.00543 -0.0089 0.0809 0.9713
3.500 0.4398 0.01275 0.00567 -0.0100 0.0605 0.9723
3.750 0.4711 0.01312 0.00599 -0.0110 0.0527 0.9735
4.000 0.5024 0.01337 0.00626 -0.0119 0.0502 0.9750
4.250 0.5320 0.01366 0.00658 -0.0125 0.0478 0.9767
4.500 0.5600 0.01398 0.00690 -0.0128 0.0454 0.9789
4.750 0.5831 0.01456 0.00749 -0.0121 0.0430 0.9819
5.000 0.6080 0.01515 0.00813 -0.0116 0.0417 0.9840
5.250 0.6392 0.01552 0.00856 -0.0126 0.0406 0.9849
5.500 0.6700 0.01592 0.00900 -0.0136 0.0390 0.9861
5.750 0.6999 0.01639 0.00950 -0.0143 0.0374 0.9875
6.000 0.7285 0.01691 0.01006 -0.0148 0.0360 0.9890
6.250 0.7551 0.01780 0.01097 -0.0149 0.0342 0.9904
6.500 0.7814 0.01901 0.01231 -0.0149 0.0328 0.9918
6.750 0.8086 0.01955 0.01296 -0.0150 0.0318 0.9936
7.000 0.8370 0.02020 0.01373 -0.0154 0.0304 0.9947
7.250 0.8659 0.02069 0.01428 -0.0161 0.0288 0.9958
7.500 0.8942 0.02125 0.01486 -0.0168 0.0275 0.9972
7.750 0.9164 0.02389 0.01774 -0.0164 0.0260 0.9981
8.000 0.9443 0.02445 0.01851 -0.0167 0.0251 0.9996
8.250 0.9635 0.02537 0.01963 -0.0153 0.0240 1.0000
8.500 0.9783 0.02606 0.02043 -0.0131 0.0230 1.0000
8.750 0.9931 0.02647 0.02089 -0.0110 0.0222 1.0000
9.000 1.0070 0.02693 0.02139 -0.0087 0.0216 1.0000
9.250 1.0171 0.02830 0.02284 -0.0060 0.0209 1.0000
9.500 1.0161 0.03120 0.02613 -0.0014 0.0203 1.0000
9.750 1.0197 0.03268 0.02792 0.0027 0.0198 1.0000
10.000 1.0159 0.03504 0.03062 0.0078 0.0193 1.0000
10.250 1.0045 0.03792 0.03386 0.0139 0.0188 1.0000
10.500 0.9867 0.04099 0.03725 0.0207 0.0185 1.0000
10.750 0.9620 0.04420 0.04075 0.0282 0.0182 1.0000
11.000 0.9220 0.04757 0.04437 0.0378 0.0183 1.0000
11.250 0.8741 0.05160 0.04862 0.0469 0.0185 1.0000
11.500 0.8222 0.05766 0.05493 0.0525 0.0188 1.0000
11.750 0.7763 0.06496 0.06241 0.0530 0.0193 1.0000
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Polar data table (+)
Polar graphs
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