USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 50,000 Max Cl/Cd: 32.37 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa51-il-50000.txt Download as CSV file: xf-usa51-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: USA 51 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4466 0.11892 0.11191 -0.0116 1.0000 0.1989
-9.500 -0.4331 0.11433 0.10732 -0.0107 1.0000 0.2082
-9.250 -0.4511 0.11358 0.10671 -0.0129 1.0000 0.2143
-9.000 -0.4355 0.10895 0.10209 -0.0116 1.0000 0.2273
-8.750 -0.4259 0.10500 0.09818 -0.0109 1.0000 0.2381
-8.500 -0.4291 0.10221 0.09548 -0.0111 1.0000 0.2472
-8.250 -0.4436 0.10090 0.09431 -0.0117 1.0000 0.2580
-8.000 -0.4312 0.09689 0.09032 -0.0104 1.0000 0.2714
-7.750 -0.4255 0.09350 0.08698 -0.0093 1.0000 0.2849
-7.500 -0.4239 0.09051 0.08407 -0.0082 1.0000 0.2991
-7.250 -0.4289 0.08806 0.08173 -0.0068 1.0000 0.3147
-7.000 -0.4387 0.08597 0.07977 -0.0049 1.0000 0.3302
-6.750 -0.4160 0.08171 0.07550 -0.0026 1.0000 0.3536
-6.500 -0.4119 0.07900 0.07284 0.0003 1.0000 0.3819
-6.250 -0.4077 0.07651 0.07042 0.0037 1.0000 0.4134
-6.000 -0.4175 0.07490 0.06895 0.0081 1.0000 0.4464
-5.750 -0.3936 0.07129 0.06533 0.0113 1.0000 0.4886
-5.500 -0.3838 0.06877 0.06286 0.0155 1.0000 0.5337
-5.000 -0.4294 0.04924 0.04219 -0.0263 1.0000 0.1736
-4.750 -0.4118 0.04464 0.03710 -0.0262 1.0000 0.1525
-4.500 -0.3941 0.04060 0.03241 -0.0255 1.0000 0.1400
-4.250 -0.3759 0.03722 0.02861 -0.0243 1.0000 0.1340
-4.000 -0.3548 0.03443 0.02482 -0.0228 1.0000 0.1285
-3.750 -0.3356 0.03201 0.02246 -0.0216 1.0000 0.1338
-3.500 -0.3141 0.03006 0.02013 -0.0203 1.0000 0.1385
-3.250 -0.2906 0.02815 0.01773 -0.0189 1.0000 0.1416
-3.000 -0.2681 0.02647 0.01597 -0.0177 1.0000 0.1496
-2.750 -0.2451 0.02514 0.01451 -0.0166 1.0000 0.1628
-2.500 -0.2191 0.02378 0.01305 -0.0156 1.0000 0.1776
-2.250 -0.1067 0.01758 0.01024 -0.0258 1.0000 1.0000
-2.000 -0.0959 0.01761 0.00979 -0.0232 1.0000 1.0000
-1.750 -0.0848 0.01769 0.00950 -0.0208 1.0000 1.0000
-1.500 -0.0730 0.01784 0.00934 -0.0186 1.0000 1.0000
-1.250 -0.0605 0.01805 0.00928 -0.0167 1.0000 1.0000
-1.000 -0.0476 0.01831 0.00931 -0.0149 1.0000 1.0000
-0.750 -0.0343 0.01863 0.00942 -0.0133 1.0000 1.0000
-0.500 -0.0211 0.01903 0.00964 -0.0120 1.0000 1.0000
-0.250 -0.0082 0.01951 0.00995 -0.0108 1.0000 1.0000
0.000 0.0039 0.02011 0.01041 -0.0097 1.0000 1.0000
0.250 0.0396 0.02108 0.01124 -0.0136 0.9900 1.0000
0.500 0.1147 0.02222 0.01219 -0.0241 0.9616 1.0000
0.750 0.1844 0.02292 0.01278 -0.0330 0.9325 1.0000
1.000 0.2485 0.02329 0.01308 -0.0403 0.9043 1.0000
1.250 0.3107 0.02338 0.01316 -0.0467 0.8778 1.0000
1.500 0.3768 0.02317 0.01297 -0.0533 0.8554 1.0000
2.000 0.4699 0.02305 0.01286 -0.0588 0.8104 1.0000
2.250 0.5076 0.02315 0.01299 -0.0599 0.7894 1.0000
2.500 0.5457 0.02324 0.01308 -0.0609 0.7713 1.0000
2.750 0.5703 0.02377 0.01363 -0.0601 0.7514 1.0000
3.000 0.5978 0.02415 0.01406 -0.0595 0.7326 1.0000
3.250 0.6287 0.02423 0.01414 -0.0588 0.7135 1.0000
3.500 0.6498 0.02464 0.01457 -0.0569 0.6916 1.0000
3.750 0.6782 0.02449 0.01437 -0.0551 0.6688 1.0000
4.000 0.6999 0.02455 0.01446 -0.0527 0.6433 1.0000
4.250 0.7217 0.02465 0.01457 -0.0505 0.6189 1.0000
4.500 0.7474 0.02468 0.01455 -0.0487 0.5974 1.0000
4.750 0.7675 0.02513 0.01510 -0.0468 0.5755 1.0000
5.000 0.7927 0.02534 0.01533 -0.0453 0.5544 1.0000
5.250 0.8132 0.02581 0.01589 -0.0433 0.5297 1.0000
5.500 0.8349 0.02623 0.01632 -0.0413 0.5017 1.0000
5.750 0.8576 0.02667 0.01666 -0.0392 0.4692 1.0000
6.000 0.8753 0.02755 0.01752 -0.0367 0.4312 1.0000
6.250 0.8963 0.02853 0.01840 -0.0348 0.3975 1.0000
6.500 0.9150 0.02979 0.01973 -0.0330 0.3712 1.0000
6.750 0.9375 0.03097 0.02093 -0.0319 0.3537 1.0000
7.000 0.9558 0.03181 0.02192 -0.0302 0.3357 1.0000
7.250 0.9731 0.03194 0.02217 -0.0282 0.3155 1.0000
7.500 0.9920 0.03154 0.02168 -0.0263 0.2958 1.0000
7.750 1.0093 0.03192 0.02225 -0.0245 0.2807 1.0000
8.000 1.0254 0.03220 0.02273 -0.0225 0.2650 1.0000
8.250 1.0391 0.03234 0.02304 -0.0202 0.2468 1.0000
8.500 1.0519 0.03250 0.02323 -0.0177 0.2272 1.0000
8.750 1.0561 0.03298 0.02375 -0.0144 0.1996 1.0000
9.000 1.0578 0.03426 0.02501 -0.0111 0.1696 1.0000
9.250 1.0587 0.03659 0.02719 -0.0078 0.1394 1.0000
9.500 1.0653 0.03944 0.02974 -0.0052 0.1170 1.0000
9.750 1.0769 0.04203 0.03233 -0.0033 0.1032 1.0000
10.000 1.0856 0.04437 0.03497 -0.0011 0.0936 1.0000
10.250 1.0975 0.04705 0.03782 0.0005 0.0873 1.0000
10.500 1.1080 0.05009 0.04118 0.0022 0.0833 1.0000
10.750 1.1191 0.05311 0.04437 0.0037 0.0803 1.0000
11.000 1.1303 0.05669 0.04805 0.0048 0.0783 1.0000
11.250 1.1225 0.06041 0.05220 0.0073 0.0778 1.0000
11.500 1.1093 0.06430 0.05641 0.0095 0.0776 1.0000
11.750 1.0934 0.06840 0.06077 0.0109 0.0777 1.0000
12.000 1.0744 0.07294 0.06554 0.0113 0.0778 1.0000
12.250 1.0547 0.07801 0.07079 0.0106 0.0782 1.0000
12.500 0.9682 0.09136 0.08452 0.0004 0.0833 1.0000
12.750 0.9223 0.10434 0.09753 -0.0084 0.0870 1.0000
13.000 0.9003 0.11410 0.10725 -0.0137 0.0890 1.0000
13.250 0.8860 0.12277 0.11588 -0.0178 0.0900 1.0000
13.500 0.8782 0.13027 0.12334 -0.0209 0.0907 1.0000
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