NACA 63A210 (naca63a210-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63A210 (naca63a210-il) Reynolds number: 100,000 Max Cl/Cd: 40.2 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca63a210-il-100000-n5.txt Download as CSV file: xf-naca63a210-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A210
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5920 0.08842 0.08390 -0.0187 1.0000 0.0280
-9.500 -0.5977 0.08179 0.07731 -0.0234 1.0000 0.0276
-9.250 -0.6082 0.07379 0.06932 -0.0306 1.0000 0.0271
-9.000 -0.6239 0.06753 0.06299 -0.0356 1.0000 0.0267
-8.750 -0.6354 0.06213 0.05743 -0.0384 0.9692 0.0264
-8.500 -0.6401 0.05676 0.05181 -0.0405 0.9478 0.0261
-8.250 -0.6399 0.05172 0.04643 -0.0416 0.9343 0.0260
-8.000 -0.6347 0.04702 0.04133 -0.0421 0.9223 0.0260
-7.750 -0.6251 0.04275 0.03660 -0.0422 0.9109 0.0262
-7.500 -0.6114 0.03929 0.03266 -0.0419 0.9022 0.0274
-7.250 -0.5943 0.03632 0.02916 -0.0414 0.8943 0.0291
-7.000 -0.5752 0.03357 0.02587 -0.0408 0.8867 0.0301
-6.750 -0.5540 0.03099 0.02282 -0.0402 0.8796 0.0305
-6.500 -0.5310 0.02887 0.02026 -0.0397 0.8728 0.0310
-6.250 -0.5083 0.02611 0.01722 -0.0393 0.8654 0.0321
-6.000 -0.4841 0.02442 0.01539 -0.0390 0.8577 0.0339
-5.750 -0.4594 0.02332 0.01416 -0.0387 0.8500 0.0371
-5.500 -0.4340 0.02205 0.01272 -0.0383 0.8428 0.0397
-5.250 -0.4091 0.02081 0.01133 -0.0376 0.8345 0.0416
-5.000 -0.3842 0.01977 0.01012 -0.0370 0.8214 0.0436
-4.750 -0.3613 0.01855 0.00887 -0.0363 0.8028 0.0476
-4.500 -0.3364 0.01785 0.00804 -0.0359 0.7826 0.0553
-4.250 -0.3123 0.01703 0.00713 -0.0353 0.7624 0.0658
-4.000 -0.2881 0.01627 0.00631 -0.0347 0.7438 0.0862
-3.750 -0.2659 0.01517 0.00554 -0.0340 0.7271 0.1571
-3.500 -0.2488 0.01353 0.00505 -0.0329 0.7121 0.4030
-3.250 -0.2278 0.01304 0.00497 -0.0313 0.6948 0.5394
-3.000 -0.2054 0.01289 0.00493 -0.0296 0.6734 0.6169
-2.750 -0.1833 0.01285 0.00493 -0.0278 0.6615 0.6785
-2.500 -0.1610 0.01285 0.00490 -0.0259 0.6558 0.7256
-2.000 -0.1120 0.01271 0.00462 -0.0236 0.6482 0.7775
-1.750 -0.0858 0.01262 0.00444 -0.0231 0.6460 0.7945
-1.500 -0.0593 0.01252 0.00424 -0.0226 0.6439 0.8103
-1.250 -0.0326 0.01243 0.00408 -0.0222 0.6420 0.8261
-1.000 -0.0055 0.01236 0.00395 -0.0218 0.6401 0.8420
-0.750 0.0225 0.01229 0.00385 -0.0216 0.6384 0.8584
-0.500 0.0516 0.01224 0.00375 -0.0217 0.6362 0.8754
-0.250 0.0826 0.01222 0.00368 -0.0222 0.6334 0.8927
0.000 0.1159 0.01221 0.00364 -0.0233 0.6303 0.9104
0.250 0.1515 0.01221 0.00360 -0.0249 0.6270 0.9282
0.500 0.1887 0.01224 0.00358 -0.0269 0.6229 0.9460
0.750 0.2271 0.01226 0.00357 -0.0293 0.6182 0.9640
1.000 0.2659 0.01229 0.00359 -0.0318 0.6141 0.9824
1.250 0.3066 0.01231 0.00362 -0.0349 0.6109 0.9993
1.500 0.3311 0.01237 0.00367 -0.0346 0.6071 1.0000
1.750 0.3555 0.01244 0.00377 -0.0343 0.6031 1.0000
2.000 0.3806 0.01254 0.00393 -0.0340 0.5989 1.0000
2.250 0.4060 0.01263 0.00409 -0.0337 0.5940 1.0000
2.500 0.4315 0.01272 0.00430 -0.0335 0.5869 1.0000
2.750 0.4568 0.01285 0.00450 -0.0332 0.5764 1.0000
3.000 0.4817 0.01309 0.00467 -0.0325 0.5660 1.0000
3.250 0.5072 0.01347 0.00483 -0.0320 0.5566 1.0000
3.500 0.5363 0.01448 0.00541 -0.0322 0.5443 1.0000
3.750 0.5645 0.01687 0.00746 -0.0328 0.5235 1.0000
4.000 0.5849 0.01933 0.01010 -0.0323 0.5004 1.0000
4.250 0.6051 0.01562 0.00716 -0.0300 0.4567 1.0000
4.500 0.6287 0.01626 0.00759 -0.0293 0.4273 1.0000
4.750 0.6521 0.01622 0.00785 -0.0284 0.3451 1.0000
5.000 0.6678 0.01729 0.00808 -0.0269 0.1913 1.0000
5.250 0.6841 0.01866 0.00884 -0.0255 0.1005 1.0000
5.500 0.7028 0.01975 0.00974 -0.0243 0.0697 1.0000
5.750 0.7226 0.02069 0.01066 -0.0232 0.0556 1.0000
6.000 0.7431 0.02155 0.01164 -0.0220 0.0480 1.0000
6.250 0.7617 0.02261 0.01276 -0.0207 0.0434 1.0000
6.500 0.7789 0.02393 0.01413 -0.0191 0.0406 1.0000
6.750 0.7990 0.02505 0.01539 -0.0178 0.0379 1.0000
7.000 0.8195 0.02616 0.01664 -0.0168 0.0346 1.0000
7.250 0.8396 0.02741 0.01794 -0.0159 0.0322 1.0000
7.500 0.8596 0.02909 0.01969 -0.0148 0.0308 1.0000
7.750 0.8797 0.03160 0.02232 -0.0139 0.0296 1.0000
8.000 0.9011 0.03357 0.02453 -0.0129 0.0290 1.0000
8.250 0.9209 0.03580 0.02709 -0.0118 0.0285 1.0000
8.500 0.9384 0.03836 0.03007 -0.0105 0.0280 1.0000
8.750 0.9526 0.04114 0.03328 -0.0090 0.0275 1.0000
9.000 0.9637 0.04388 0.03646 -0.0072 0.0265 1.0000
9.250 0.9717 0.04666 0.03963 -0.0055 0.0252 1.0000
9.500 0.9762 0.04968 0.04302 -0.0037 0.0243 1.0000
9.750 0.9751 0.05328 0.04699 -0.0018 0.0241 1.0000
10.000 0.9685 0.05707 0.05113 0.0001 0.0239 1.0000
10.250 0.9554 0.06085 0.05519 0.0021 0.0239 1.0000
10.500 0.9378 0.06496 0.05955 0.0031 0.0239 1.0000
10.750 0.9171 0.06991 0.06472 0.0022 0.0241 1.0000
11.000 0.8940 0.07624 0.07123 -0.0012 0.0243 1.0000
11.250 0.8685 0.08454 0.07968 -0.0075 0.0247 1.0000
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Polar data table (+)
Polar graphs
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