NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Reynolds number: 50,000 Max Cl/Cd: 25.21 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rcsc2-il-50000-n5.txt Download as CSV file: xf-rcsc2-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-SC2 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6014 0.10242 0.09414 -0.0297 1.0000 0.0624
-10.250 -0.6024 0.09781 0.08956 -0.0314 1.0000 0.0620
-10.000 -0.6078 0.09228 0.08405 -0.0343 1.0000 0.0616
-9.750 -0.6185 0.08639 0.07817 -0.0377 1.0000 0.0612
-9.500 -0.6344 0.08111 0.07289 -0.0398 1.0000 0.0608
-9.250 -0.6539 0.07670 0.06845 -0.0399 1.0000 0.0604
-9.000 -0.6730 0.07271 0.06440 -0.0386 1.0000 0.0602
-8.750 -0.6889 0.06876 0.06032 -0.0369 1.0000 0.0601
-8.500 -0.7018 0.06497 0.05635 -0.0348 1.0000 0.0601
-8.250 -0.7113 0.06132 0.05245 -0.0324 1.0000 0.0603
-8.000 -0.7172 0.05778 0.04862 -0.0298 1.0000 0.0606
-7.750 -0.7193 0.05441 0.04491 -0.0272 1.0000 0.0610
-7.500 -0.7175 0.05122 0.04136 -0.0246 1.0000 0.0613
-7.250 -0.7122 0.04822 0.03797 -0.0221 1.0000 0.0616
-7.000 -0.7039 0.04539 0.03474 -0.0197 1.0000 0.0619
-6.750 -0.6930 0.04275 0.03166 -0.0174 1.0000 0.0624
-6.500 -0.6786 0.04042 0.02911 -0.0156 1.0000 0.0631
-6.250 -0.6624 0.03854 0.02709 -0.0141 1.0000 0.0645
-6.000 -0.6451 0.03692 0.02533 -0.0125 1.0000 0.0666
-5.750 -0.6267 0.03527 0.02344 -0.0110 1.0000 0.0691
-5.500 -0.6066 0.03357 0.02142 -0.0096 1.0000 0.0716
-5.250 -0.5847 0.03196 0.01947 -0.0082 1.0000 0.0737
-5.000 -0.5632 0.03056 0.01812 -0.0072 1.0000 0.0765
-4.750 -0.5414 0.02944 0.01694 -0.0061 1.0000 0.0811
-4.500 -0.5179 0.02832 0.01563 -0.0049 1.0000 0.0864
-4.250 -0.4957 0.02724 0.01462 -0.0038 1.0000 0.0918
-4.000 -0.4728 0.02631 0.01360 -0.0027 1.0000 0.1005
-3.750 -0.4507 0.02541 0.01272 -0.0016 1.0000 0.1118
-3.500 -0.4288 0.02452 0.01184 -0.0005 1.0000 0.1283
-3.250 -0.4078 0.02355 0.01108 0.0005 1.0000 0.1543
-3.000 -0.3880 0.02246 0.01037 0.0016 1.0000 0.2038
-2.750 -0.3716 0.02098 0.00970 0.0032 1.0000 0.3162
-2.500 -0.3603 0.01933 0.00962 0.0070 1.0000 0.5859
-2.250 -0.3249 0.01942 0.01036 0.0089 1.0000 0.8135
-2.000 -0.2612 0.02024 0.01098 0.0046 1.0000 0.9235
-1.750 -0.1326 0.02091 0.01113 -0.0138 1.0000 0.9887
-1.500 -0.0929 0.02072 0.01075 -0.0170 1.0000 1.0000
-1.250 -0.0852 0.02052 0.01047 -0.0140 1.0000 1.0000
-1.000 -0.0771 0.02037 0.01023 -0.0110 1.0000 1.0000
-0.750 -0.0684 0.02025 0.01004 -0.0081 1.0000 1.0000
-0.500 -0.0592 0.02017 0.00989 -0.0053 1.0000 1.0000
-0.250 -0.0492 0.02013 0.00979 -0.0026 1.0000 1.0000
0.000 -0.0385 0.02013 0.00973 0.0001 1.0000 1.0000
0.250 -0.0270 0.02016 0.00971 0.0025 1.0000 1.0000
0.500 -0.0149 0.02023 0.00974 0.0049 1.0000 1.0000
0.750 -0.0021 0.02034 0.00982 0.0071 1.0000 1.0000
1.000 0.0115 0.02049 0.00995 0.0091 1.0000 1.0000
1.250 0.0259 0.02069 0.01013 0.0109 1.0000 1.0000
1.500 0.0410 0.02093 0.01038 0.0126 1.0000 1.0000
1.750 0.0800 0.02137 0.01087 0.0095 0.9916 1.0000
2.000 0.1463 0.02190 0.01152 0.0015 0.9725 1.0000
2.250 0.2084 0.02206 0.01186 -0.0052 0.9487 1.0000
2.500 0.2617 0.02200 0.01198 -0.0098 0.9241 1.0000
2.750 0.3134 0.02170 0.01191 -0.0137 0.8962 1.0000
3.000 0.3616 0.02115 0.01159 -0.0163 0.8607 1.0000
3.250 0.4001 0.02044 0.01109 -0.0165 0.8109 1.0000
3.500 0.4469 0.01962 0.01034 -0.0174 0.7174 1.0000
3.750 0.5001 0.01984 0.00939 -0.0189 0.4677 1.0000
4.000 0.5106 0.02109 0.00981 -0.0155 0.3459 1.0000
4.250 0.5246 0.02227 0.01046 -0.0131 0.2720 1.0000
4.500 0.5414 0.02335 0.01120 -0.0113 0.2206 1.0000
4.750 0.5601 0.02438 0.01201 -0.0098 0.1833 1.0000
5.000 0.5797 0.02539 0.01281 -0.0085 0.1543 1.0000
5.250 0.6008 0.02634 0.01368 -0.0073 0.1318 1.0000
5.500 0.6231 0.02737 0.01465 -0.0064 0.1157 1.0000
5.750 0.6464 0.02847 0.01576 -0.0054 0.1034 1.0000
6.000 0.6698 0.02966 0.01695 -0.0045 0.0940 1.0000
6.250 0.6927 0.03089 0.01817 -0.0037 0.0873 1.0000
6.500 0.7166 0.03233 0.01976 -0.0028 0.0815 1.0000
6.750 0.7383 0.03366 0.02117 -0.0018 0.0766 1.0000
7.000 0.7600 0.03524 0.02270 -0.0009 0.0735 1.0000
7.250 0.7816 0.03723 0.02508 0.0003 0.0710 1.0000
7.500 0.8004 0.03922 0.02740 0.0017 0.0685 1.0000
7.750 0.8176 0.04107 0.02946 0.0032 0.0661 1.0000
8.000 0.8340 0.04282 0.03130 0.0045 0.0640 1.0000
8.250 0.8484 0.04505 0.03368 0.0061 0.0625 1.0000
8.500 0.8574 0.04792 0.03706 0.0084 0.0616 1.0000
8.750 0.8634 0.05102 0.04059 0.0107 0.0610 1.0000
9.000 0.8659 0.05432 0.04430 0.0132 0.0606 1.0000
9.250 0.8643 0.05782 0.04817 0.0157 0.0603 1.0000
9.500 0.8585 0.06152 0.05219 0.0181 0.0602 1.0000
9.750 0.8482 0.06539 0.05635 0.0204 0.0602 1.0000
10.000 0.8330 0.06939 0.06058 0.0226 0.0604 1.0000
10.250 0.8118 0.07353 0.06489 0.0245 0.0606 1.0000
10.500 0.7882 0.07847 0.06998 0.0247 0.0609 1.0000
10.750 0.7642 0.08447 0.07608 0.0227 0.0613 1.0000
11.000 0.7430 0.09151 0.08318 0.0187 0.0618 1.0000
11.250 0.7285 0.09892 0.09061 0.0140 0.0623 1.0000
11.500 0.7211 0.10573 0.09741 0.0103 0.0627 1.0000
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Polar data table (+)
Polar graphs
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