NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il) Reynolds number: 50,000 Max Cl/Cd: 27.68 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc410-il-50000-n5.txt Download as CSV file: xf-rc410-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(4)-10 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5361 0.10647 0.09949 -0.0014 1.0000 0.0641
-9.250 -0.5323 0.10222 0.09529 -0.0033 1.0000 0.0638
-9.000 -0.5311 0.09765 0.09076 -0.0062 1.0000 0.0637
-8.750 -0.5315 0.09276 0.08592 -0.0099 1.0000 0.0636
-8.500 -0.5337 0.08774 0.08094 -0.0139 1.0000 0.0635
-8.250 -0.5360 0.08293 0.07614 -0.0170 1.0000 0.0634
-8.000 -0.5364 0.07804 0.07122 -0.0199 1.0000 0.0633
-7.750 -0.5355 0.07321 0.06631 -0.0223 1.0000 0.0633
-7.500 -0.5326 0.06848 0.06147 -0.0242 1.0000 0.0632
-7.250 -0.5278 0.06385 0.05669 -0.0256 1.0000 0.0631
-7.000 -0.5216 0.05925 0.05188 -0.0266 1.0000 0.0631
-6.750 -0.5143 0.05469 0.04704 -0.0271 1.0000 0.0634
-6.500 -0.5070 0.05005 0.04200 -0.0270 1.0000 0.0640
-6.250 -0.4931 0.04804 0.03996 -0.0263 1.0000 0.0658
-6.000 -0.4813 0.04632 0.03817 -0.0251 1.0000 0.0680
-5.750 -0.4756 0.04416 0.03577 -0.0230 1.0000 0.0701
-5.500 -0.4715 0.04174 0.03302 -0.0203 1.0000 0.0720
-5.000 -0.4219 0.03651 0.02702 -0.0219 0.9836 0.0789
-4.750 -0.3901 0.03415 0.02410 -0.0234 0.9740 0.0861
-4.500 -0.3571 0.03283 0.02264 -0.0251 0.9657 0.0932
-4.250 -0.3227 0.03111 0.02052 -0.0267 0.9582 0.1031
-4.000 -0.2899 0.02994 0.01898 -0.0279 0.9503 0.1144
-3.750 -0.2563 0.02891 0.01789 -0.0294 0.9432 0.1255
-3.500 -0.2242 0.02791 0.01671 -0.0304 0.9360 0.1371
-3.250 -0.1905 0.02699 0.01557 -0.0316 0.9291 0.1489
-3.000 -0.1563 0.02622 0.01462 -0.0327 0.9229 0.1604
-2.750 -0.1242 0.02547 0.01389 -0.0336 0.9160 0.1712
-2.500 -0.0877 0.02480 0.01312 -0.0351 0.9108 0.1816
-2.250 -0.0567 0.02437 0.01258 -0.0357 0.9040 0.1934
-2.000 -0.0260 0.02384 0.01209 -0.0363 0.8977 0.2039
-1.750 0.0031 0.02344 0.01170 -0.0366 0.8919 0.2164
-1.500 0.0280 0.02315 0.01143 -0.0363 0.8847 0.2303
-1.250 0.0562 0.02280 0.01115 -0.0364 0.8791 0.2476
-1.000 0.0794 0.02245 0.01100 -0.0359 0.8721 0.2766
-0.750 0.1390 0.02017 0.01124 -0.0400 0.8711 0.8889
-0.500 0.2228 0.02018 0.01097 -0.0501 0.8684 1.0000
-0.250 0.2464 0.02036 0.01096 -0.0491 0.8568 1.0000
0.000 0.2691 0.02050 0.01093 -0.0476 0.8426 1.0000
0.250 0.2901 0.02059 0.01087 -0.0455 0.8262 1.0000
0.500 0.3102 0.02067 0.01081 -0.0433 0.8091 1.0000
0.750 0.3304 0.02074 0.01075 -0.0410 0.7928 1.0000
1.000 0.3508 0.02079 0.01071 -0.0388 0.7769 1.0000
1.250 0.3710 0.02082 0.01064 -0.0366 0.7601 1.0000
1.500 0.3911 0.02079 0.01054 -0.0343 0.7419 1.0000
1.750 0.4115 0.02073 0.01042 -0.0320 0.7234 1.0000
2.000 0.4324 0.02068 0.01032 -0.0299 0.7055 1.0000
2.250 0.4536 0.02060 0.01019 -0.0278 0.6868 1.0000
2.500 0.4743 0.02054 0.01013 -0.0258 0.6634 1.0000
2.750 0.4953 0.02037 0.00993 -0.0236 0.6374 1.0000
3.000 0.5158 0.02029 0.00984 -0.0216 0.6012 1.0000
3.250 0.5361 0.02019 0.00967 -0.0193 0.5498 1.0000
3.500 0.5556 0.02015 0.00928 -0.0166 0.4631 1.0000
3.750 0.5727 0.02069 0.00901 -0.0139 0.3834 1.0000
4.000 0.5906 0.02163 0.00945 -0.0124 0.3428 1.0000
4.250 0.6100 0.02252 0.01005 -0.0112 0.3185 1.0000
4.500 0.6304 0.02333 0.01069 -0.0101 0.3014 1.0000
4.750 0.6514 0.02412 0.01132 -0.0090 0.2876 1.0000
5.000 0.6732 0.02489 0.01197 -0.0080 0.2754 1.0000
5.250 0.6959 0.02561 0.01268 -0.0072 0.2646 1.0000
5.500 0.7192 0.02638 0.01335 -0.0064 0.2562 1.0000
5.750 0.7430 0.02712 0.01412 -0.0057 0.2474 1.0000
6.000 0.7671 0.02792 0.01487 -0.0050 0.2398 1.0000
6.250 0.7911 0.02870 0.01574 -0.0044 0.2318 1.0000
6.500 0.8162 0.02958 0.01650 -0.0038 0.2260 1.0000
6.750 0.8400 0.03050 0.01760 -0.0032 0.2194 1.0000
7.000 0.8638 0.03142 0.01857 -0.0027 0.2131 1.0000
7.250 0.8876 0.03244 0.01961 -0.0022 0.2075 1.0000
7.500 0.9103 0.03357 0.02095 -0.0016 0.2018 1.0000
7.750 0.9336 0.03467 0.02210 -0.0011 0.1970 1.0000
8.000 0.9563 0.03588 0.02338 -0.0006 0.1923 1.0000
8.250 0.9757 0.03726 0.02508 0.0002 0.1867 1.0000
8.500 0.9968 0.03854 0.02647 0.0008 0.1821 1.0000
8.750 1.0198 0.03984 0.02777 0.0012 0.1785 1.0000
9.000 1.0347 0.04174 0.03010 0.0021 0.1739 1.0000
9.250 1.0508 0.04344 0.03206 0.0030 0.1691 1.0000
9.500 1.0698 0.04484 0.03355 0.0037 0.1652 1.0000
9.750 1.0874 0.04663 0.03547 0.0043 0.1621 1.0000
10.000 1.0920 0.04941 0.03876 0.0056 0.1582 1.0000
10.250 1.0995 0.05177 0.04142 0.0067 0.1543 1.0000
10.500 1.1124 0.05352 0.04330 0.0076 0.1508 1.0000
10.750 1.1315 0.05497 0.04473 0.0082 0.1479 1.0000
11.000 1.1144 0.05931 0.04961 0.0098 0.1456 1.0000
11.250 1.0916 0.06392 0.05459 0.0111 0.1435 1.0000
11.500 1.0567 0.06926 0.06017 0.0119 0.1423 1.0000
11.750 0.9983 0.07854 0.06965 0.0086 0.1418 1.0000
12.000 0.9011 0.09848 0.08963 -0.0039 0.1410 1.0000
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Polar data table (+)
Polar graphs
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