NACA 65-210 (naca65210-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 65-210 (naca65210-il) Reynolds number: 500,000 Max Cl/Cd: 62.73 at α=2.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca65210-il-500000-n5.txt Download as CSV file: xf-naca65210-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65-210
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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5774 0.08645 0.08418 -0.0285 1.0000 0.0070
-10.250 -0.5903 0.07868 0.07644 -0.0336 1.0000 0.0070
-10.000 -0.6148 0.06765 0.06542 -0.0435 1.0000 0.0069
-9.750 -0.6693 0.05724 0.05485 -0.0482 1.0000 0.0065
-9.500 -0.6786 0.05442 0.05194 -0.0472 1.0000 0.0067
-9.250 -0.6910 0.04966 0.04701 -0.0463 1.0000 0.0068
-9.000 -0.7059 0.04347 0.04052 -0.0445 1.0000 0.0069
-8.750 -0.7158 0.03755 0.03427 -0.0418 1.0000 0.0072
-8.500 -0.7238 0.02392 0.01927 -0.0410 0.9857 0.0084
-8.250 -0.6982 0.02123 0.01611 -0.0418 0.9780 0.0088
-8.000 -0.6706 0.01952 0.01417 -0.0428 0.9708 0.0094
-7.750 -0.6397 0.01893 0.01350 -0.0442 0.9643 0.0099
-7.500 -0.6101 0.01868 0.01320 -0.0451 0.9560 0.0105
-7.250 -0.5821 0.01785 0.01222 -0.0456 0.9478 0.0114
-7.000 -0.5560 0.01687 0.01105 -0.0457 0.9386 0.0121
-6.750 -0.5309 0.01605 0.01004 -0.0454 0.9292 0.0126
-6.500 -0.5074 0.01499 0.00883 -0.0449 0.9200 0.0133
-6.250 -0.4835 0.01433 0.00810 -0.0444 0.9106 0.0142
-6.000 -0.4582 0.01401 0.00774 -0.0441 0.9017 0.0150
-5.750 -0.4333 0.01359 0.00724 -0.0438 0.8934 0.0159
-5.500 -0.4088 0.01303 0.00658 -0.0433 0.8847 0.0167
-5.250 -0.3841 0.01252 0.00597 -0.0429 0.8768 0.0174
-5.000 -0.3589 0.01209 0.00545 -0.0425 0.8689 0.0181
-4.750 -0.3329 0.01182 0.00512 -0.0423 0.8615 0.0188
-4.500 -0.3092 0.01108 0.00429 -0.0418 0.8535 0.0203
-4.250 -0.2834 0.01069 0.00384 -0.0415 0.8461 0.0216
-4.000 -0.2572 0.01040 0.00349 -0.0414 0.8388 0.0230
-3.750 -0.2306 0.01014 0.00317 -0.0412 0.8321 0.0248
-3.500 -0.2038 0.00992 0.00289 -0.0411 0.8250 0.0266
-3.250 -0.1770 0.00970 0.00260 -0.0411 0.8181 0.0292
-3.000 -0.1501 0.00946 0.00234 -0.0410 0.8110 0.0346
-2.750 -0.1229 0.00930 0.00214 -0.0410 0.8047 0.0398
-2.500 -0.0959 0.00904 0.00195 -0.0410 0.7981 0.0626
-2.250 -0.0708 0.00836 0.00168 -0.0410 0.7919 0.1893
-2.000 -0.0469 0.00733 0.00139 -0.0411 0.7850 0.4008
-1.500 0.0034 0.00644 0.00136 -0.0406 0.7725 0.6636
-1.250 0.0308 0.00640 0.00136 -0.0404 0.7666 0.6889
-1.000 0.0587 0.00639 0.00135 -0.0405 0.7606 0.7019
-0.750 0.0865 0.00638 0.00135 -0.0404 0.7544 0.7170
-0.500 0.1144 0.00639 0.00134 -0.0405 0.7488 0.7265
-0.250 0.1427 0.00639 0.00133 -0.0406 0.7426 0.7330
0.000 0.1708 0.00641 0.00133 -0.0407 0.7369 0.7391
0.250 0.1990 0.00641 0.00135 -0.0408 0.7310 0.7453
0.500 0.2272 0.00644 0.00136 -0.0409 0.7250 0.7517
0.750 0.2552 0.00645 0.00139 -0.0410 0.7193 0.7577
1.000 0.2834 0.00647 0.00143 -0.0411 0.7131 0.7644
1.250 0.3112 0.00650 0.00148 -0.0411 0.7071 0.7705
1.500 0.3387 0.00654 0.00152 -0.0411 0.6939 0.7772
1.750 0.3659 0.00659 0.00155 -0.0409 0.6764 0.7834
2.000 0.3927 0.00665 0.00161 -0.0407 0.6553 0.7904
2.250 0.4186 0.00679 0.00164 -0.0403 0.6208 0.7973
2.500 0.4429 0.00706 0.00171 -0.0396 0.5574 0.8043
2.750 0.4617 0.00797 0.00196 -0.0383 0.3950 0.8115
3.000 0.4807 0.00903 0.00241 -0.0372 0.2400 0.8194
3.250 0.5012 0.00995 0.00282 -0.0364 0.1140 0.8273
3.500 0.5240 0.01056 0.00317 -0.0358 0.0486 0.8355
3.750 0.5491 0.01086 0.00344 -0.0354 0.0360 0.8430
4.000 0.5748 0.01110 0.00373 -0.0351 0.0308 0.8512
4.250 0.5996 0.01137 0.00403 -0.0346 0.0266 0.8594
4.500 0.6237 0.01175 0.00447 -0.0340 0.0230 0.8683
4.750 0.6483 0.01200 0.00479 -0.0335 0.0218 0.8776
5.000 0.6722 0.01231 0.00519 -0.0328 0.0202 0.8867
5.250 0.6956 0.01266 0.00560 -0.0321 0.0189 0.8969
5.500 0.7183 0.01303 0.00603 -0.0312 0.0177 0.9082
5.750 0.7393 0.01352 0.00659 -0.0300 0.0167 0.9205
6.000 0.7575 0.01433 0.00750 -0.0284 0.0157 0.9347
6.250 0.7794 0.01478 0.00805 -0.0274 0.0153 0.9508
6.500 0.8063 0.01527 0.00862 -0.0276 0.0147 0.9701
6.750 0.8333 0.01596 0.00938 -0.0279 0.0142 1.0000
7.000 0.8561 0.01671 0.01021 -0.0273 0.0138 1.0000
7.250 0.8788 0.01750 0.01107 -0.0268 0.0133 1.0000
7.500 0.9017 0.01825 0.01189 -0.0263 0.0128 1.0000
7.750 0.9252 0.01881 0.01252 -0.0259 0.0122 1.0000
8.000 0.9484 0.01933 0.01307 -0.0256 0.0116 1.0000
8.250 0.9697 0.02028 0.01409 -0.0249 0.0111 1.0000
8.500 0.9888 0.02205 0.01602 -0.0240 0.0106 1.0000
8.750 1.0105 0.02301 0.01714 -0.0233 0.0103 1.0000
9.000 1.0311 0.02420 0.01850 -0.0226 0.0099 1.0000
9.250 1.0510 0.02541 0.01991 -0.0218 0.0094 1.0000
9.500 1.0700 0.02655 0.02122 -0.0209 0.0089 1.0000
9.750 1.0884 0.02753 0.02234 -0.0200 0.0084 1.0000
10.000 1.1063 0.02829 0.02319 -0.0191 0.0081 1.0000
10.250 1.1224 0.02929 0.02432 -0.0179 0.0078 1.0000
10.500 1.1371 0.03028 0.02541 -0.0167 0.0076 1.0000
10.750 1.1476 0.03171 0.02699 -0.0150 0.0073 1.0000
11.000 1.1461 0.03448 0.03007 -0.0119 0.0071 1.0000
11.250 1.1438 0.03708 0.03298 -0.0089 0.0069 1.0000
11.500 1.1321 0.04067 0.03693 -0.0056 0.0068 1.0000
11.750 1.1127 0.04507 0.04170 -0.0026 0.0066 1.0000
12.000 1.0780 0.05148 0.04851 -0.0004 0.0065 1.0000
12.250 1.0363 0.05944 0.05682 -0.0004 0.0064 1.0000
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Polar data table (+)
Polar graphs
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