NASA SC(2)-0404 AIRFOIL (sc20404-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA SC(2)-0404 AIRFOIL (sc20404-il) Reynolds number: 100,000 Max Cl/Cd: 25.99 at α=1.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20404-il-100000-n5.txt Download as CSV file: xf-sc20404-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0404 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6673 0.11671 0.11170 0.0232 1.0000 0.0301
-9.250 -0.6644 0.11267 0.10769 0.0212 1.0000 0.0303
-9.000 -0.6615 0.10859 0.10364 0.0191 1.0000 0.0304
-8.000 -0.5547 0.07829 0.07369 0.0001 1.0000 0.0189
-7.500 -0.6373 0.08004 0.07526 -0.0027 1.0000 0.0193
-7.250 -0.6287 0.07467 0.06985 -0.0075 1.0000 0.0184
-7.000 -0.6175 0.06914 0.06424 -0.0123 1.0000 0.0176
-6.750 -0.6036 0.06353 0.05852 -0.0169 1.0000 0.0168
-6.500 -0.5870 0.05794 0.05274 -0.0210 1.0000 0.0161
-6.250 -0.5676 0.05242 0.04696 -0.0244 1.0000 0.0154
-6.000 -0.5455 0.04712 0.04134 -0.0273 1.0000 0.0151
-5.750 -0.5210 0.04253 0.03637 -0.0294 1.0000 0.0167
-5.500 -0.4933 0.03823 0.03156 -0.0310 1.0000 0.0192
-5.250 -0.4646 0.03432 0.02708 -0.0321 1.0000 0.0204
-5.000 -0.4358 0.03012 0.02213 -0.0330 1.0000 0.0204
-4.750 -0.4063 0.02672 0.01814 -0.0335 1.0000 0.0206
-4.500 -0.3772 0.02395 0.01484 -0.0336 1.0000 0.0210
-4.250 -0.3487 0.02153 0.01203 -0.0334 1.0000 0.0217
-4.000 -0.3217 0.01892 0.00921 -0.0333 1.0000 0.0239
-3.750 -0.2947 0.01802 0.00815 -0.0332 1.0000 0.0325
-3.500 -0.2677 0.01639 0.00642 -0.0329 1.0000 0.0349
-3.250 -0.2400 0.01519 0.00512 -0.0330 1.0000 0.0400
-3.000 -0.2116 0.01423 0.00404 -0.0331 1.0000 0.0534
-2.750 -0.1834 0.01206 0.00317 -0.0345 1.0000 0.3216
-2.500 -0.1674 0.01078 0.00326 -0.0313 1.0000 0.7184
-2.250 -0.1526 0.01061 0.00325 -0.0272 1.0000 0.8070
-2.000 -0.1425 0.01039 0.00310 -0.0220 1.0000 0.8718
-1.750 -0.1290 0.01005 0.00277 -0.0179 1.0000 0.9231
-1.500 -0.1008 0.00980 0.00241 -0.0178 1.0000 0.9594
-1.250 -0.0703 0.00961 0.00212 -0.0186 1.0000 1.0000
-1.000 -0.0412 0.00958 0.00198 -0.0194 1.0000 1.0000
-0.750 -0.0122 0.00957 0.00186 -0.0201 1.0000 1.0000
-0.500 0.0166 0.00958 0.00180 -0.0207 1.0000 1.0000
-0.250 0.0451 0.00960 0.00179 -0.0212 1.0000 1.0000
0.000 0.0734 0.00963 0.00183 -0.0216 1.0000 1.0000
0.250 0.1014 0.00967 0.00191 -0.0219 1.0000 1.0000
0.500 0.1292 0.00973 0.00206 -0.0222 1.0000 1.0000
0.750 0.1569 0.00980 0.00222 -0.0224 1.0000 1.0000
1.000 0.1844 0.00988 0.00243 -0.0226 1.0000 1.0000
1.250 0.2117 0.00998 0.00270 -0.0227 1.0000 1.0000
1.500 0.2390 0.01009 0.00303 -0.0227 1.0000 1.0000
1.750 0.3111 0.01197 0.00311 -0.0297 0.3471 1.0000
2.000 0.3346 0.01418 0.00402 -0.0297 0.0547 1.0000
2.250 0.3620 0.01514 0.00507 -0.0294 0.0396 1.0000
2.500 0.3887 0.01622 0.00627 -0.0291 0.0344 1.0000
2.750 0.4148 0.01784 0.00798 -0.0286 0.0321 1.0000
3.000 0.4426 0.01944 0.00979 -0.0281 0.0295 1.0000
3.250 0.4699 0.02067 0.01117 -0.0280 0.0221 1.0000
3.750 0.5252 0.02645 0.01777 -0.0270 0.0201 1.0000
4.000 0.5525 0.02950 0.02139 -0.0264 0.0198 1.0000
4.250 0.5787 0.03310 0.02557 -0.0258 0.0196 1.0000
4.500 0.6031 0.03736 0.03053 -0.0253 0.0197 1.0000
4.750 0.6289 0.04103 0.03473 -0.0249 0.0185 1.0000
5.000 0.6535 0.04528 0.03946 -0.0248 0.0163 1.0000
5.250 0.6754 0.04978 0.04433 -0.0254 0.0148 1.0000
5.500 0.6952 0.05480 0.04966 -0.0263 0.0149 1.0000
5.750 0.7132 0.06007 0.05519 -0.0278 0.0155 1.0000
6.000 0.7290 0.06541 0.06074 -0.0298 0.0162 1.0000
6.250 0.7424 0.07078 0.06627 -0.0323 0.0171 1.0000
6.500 0.7531 0.07610 0.07170 -0.0349 0.0180 1.0000
6.750 0.7574 0.08121 0.07684 -0.0360 0.0195 1.0000
7.250 0.7793 0.09324 0.08907 -0.0478 0.0263 1.0000
7.750 0.7823 0.10305 0.09883 -0.0557 0.0278 1.0000
8.000 0.7832 0.10765 0.10340 -0.0587 0.0281 1.0000
8.250 0.7844 0.11214 0.10785 -0.0613 0.0283 1.0000
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Polar data table (+)
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