NASA SC(2)-0402 AIRFOIL (sc20402-il) Xfoil prediction polar at RE=500,000 Ncrit=0
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Airfoil: NASA SC(2)-0402 AIRFOIL (sc20402-il) Reynolds number: 500,000 Max Cl/Cd: 64.59 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20402-il-500000.txt Download as CSV file: xf-sc20402-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0402 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.7221 0.13815 0.13593 0.0467 1.0000 0.0026
-10.500 -0.7184 0.13430 0.13209 0.0455 1.0000 0.0027
-10.250 -0.7147 0.13048 0.12828 0.0441 1.0000 0.0028
-10.000 -0.7110 0.12666 0.12447 0.0427 1.0000 0.0030
-9.750 -0.7072 0.12285 0.12067 0.0412 1.0000 0.0032
-9.500 -0.7033 0.11903 0.11685 0.0396 1.0000 0.0035
-9.250 -0.6994 0.11521 0.11305 0.0380 1.0000 0.0039
-9.000 -0.6953 0.11138 0.10923 0.0362 1.0000 0.0043
-8.750 -0.6906 0.10759 0.10546 0.0344 1.0000 0.0049
-8.500 -0.6833 0.10414 0.10203 0.0321 1.0000 0.0056
-5.750 -0.5137 0.05476 0.05188 -0.0161 1.0000 0.0065
-5.500 -0.4922 0.04769 0.04463 -0.0201 1.0000 0.0067
-5.250 -0.4689 0.04169 0.03840 -0.0233 1.0000 0.0071
-5.000 -0.4448 0.03727 0.03378 -0.0256 1.0000 0.0077
-4.750 -0.4186 0.03367 0.02998 -0.0273 1.0000 0.0086
-4.500 -0.3903 0.03032 0.02639 -0.0287 1.0000 0.0097
-3.500 -0.2682 0.01910 0.01388 -0.0313 1.0000 0.0193
-3.250 -0.2393 0.01765 0.01222 -0.0316 1.0000 0.0261
-3.000 -0.2094 0.01538 0.00966 -0.0321 1.0000 0.0329
-2.750 -0.1802 0.01478 0.00887 -0.0319 1.0000 0.0408
-2.000 -0.0920 0.00974 0.00342 -0.0306 1.0000 0.0180
-1.750 -0.0600 0.00800 0.00157 -0.0309 1.0000 0.0089
-1.500 -0.0310 0.00750 0.00094 -0.0309 1.0000 0.0084
-1.250 -0.0005 0.00650 0.00047 -0.0318 1.0000 0.2319
-1.000 0.0061 0.00411 0.00051 -0.0269 1.0000 0.9894
-0.750 0.0335 0.00411 0.00046 -0.0269 1.0000 1.0000
-0.500 0.0614 0.00411 0.00046 -0.0270 1.0000 1.0000
-0.250 0.0891 0.00413 0.00048 -0.0271 1.0000 1.0000
0.000 0.1340 0.00493 0.00048 -0.0304 0.7002 1.0000
0.250 0.1554 0.00762 0.00092 -0.0301 0.0087 1.0000
0.500 0.1839 0.00805 0.00150 -0.0300 0.0085 1.0000
1.500 0.2993 0.01355 0.00745 -0.0269 0.0460 1.0000
1.750 0.3274 0.01507 0.00917 -0.0267 0.0354 1.0000
2.000 0.3383 0.00601 0.00061 -0.0242 0.0288 1.0000
2.250 0.3688 0.00571 0.00059 -0.0234 0.0222 1.0000
2.500 0.3913 0.00755 0.00259 -0.0236 0.0172 1.0000
2.750 0.4145 0.01066 0.00620 -0.0229 0.0161 1.0000
3.000 0.4452 0.01065 0.00644 -0.0217 0.0130 1.0000
3.250 0.4706 0.01277 0.00885 -0.0211 0.0109 1.0000
3.500 0.4947 0.01541 0.01174 -0.0208 0.0095 1.0000
3.750 0.5176 0.01848 0.01505 -0.0207 0.0085 1.0000
4.000 0.5384 0.02204 0.01880 -0.0210 0.0077 1.0000
4.250 0.5569 0.02638 0.02333 -0.0215 0.0072 1.0000
4.500 0.5766 0.03108 0.02850 -0.0220 0.0071 1.0000
4.750 0.5944 0.03631 0.03395 -0.0228 0.0069 1.0000
5.000 0.6085 0.04262 0.04042 -0.0242 0.0067 1.0000
5.250 0.6188 0.05006 0.04800 -0.0261 0.0065 1.0000
5.500 0.6901 0.06949 0.06733 -0.0342 0.0064 1.0000
5.750 0.7084 0.07439 0.07234 -0.0370 0.0064 1.0000
6.000 0.7254 0.07937 0.07740 -0.0402 0.0064 1.0000
6.250 0.7410 0.08439 0.08248 -0.0439 0.0064 1.0000
6.500 0.7554 0.08940 0.08754 -0.0481 0.0063 1.0000
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Polar data table (+)
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