NASA SC(2)-0010 AIRFOIL (sc20010-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-0010 AIRFOIL (sc20010-il) Reynolds number: 500,000 Max Cl/Cd: 52.04 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20010-il-500000.txt Download as CSV file: xf-sc20010-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0010 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
------ -------- --------- --------- -------- -------- --------
-12.250 -1.0785 0.04676 0.04311 -0.0260 1.0000 0.0200
-12.000 -1.0927 0.04361 0.03974 -0.0237 1.0000 0.0201
-11.750 -1.0981 0.04142 0.03739 -0.0210 1.0000 0.0202
-11.500 -1.0933 0.03961 0.03544 -0.0193 1.0000 0.0205
-11.250 -1.0879 0.03752 0.03317 -0.0175 1.0000 0.0208
-11.000 -1.0825 0.03502 0.03041 -0.0157 1.0000 0.0210
-10.750 -1.0751 0.03238 0.02747 -0.0139 1.0000 0.0212
-10.500 -1.0634 0.03011 0.02492 -0.0124 1.0000 0.0214
-10.250 -1.0485 0.02815 0.02271 -0.0111 1.0000 0.0217
-10.000 -1.0314 0.02642 0.02076 -0.0100 1.0000 0.0220
-9.750 -1.0124 0.02490 0.01903 -0.0090 1.0000 0.0223
-9.500 -0.9921 0.02357 0.01752 -0.0082 1.0000 0.0227
-9.250 -0.9705 0.02244 0.01622 -0.0074 1.0000 0.0230
-9.000 -0.9482 0.02143 0.01507 -0.0067 1.0000 0.0233
-8.750 -0.9285 0.01968 0.01321 -0.0058 1.0000 0.0241
-8.500 -0.9054 0.01892 0.01244 -0.0053 1.0000 0.0248
-8.250 -0.8815 0.01829 0.01177 -0.0049 1.0000 0.0256
-8.000 -0.8576 0.01760 0.01102 -0.0043 1.0000 0.0265
-7.750 -0.8336 0.01690 0.01025 -0.0038 1.0000 0.0273
-7.500 -0.8090 0.01632 0.00958 -0.0033 1.0000 0.0281
-7.250 -0.7870 0.01529 0.00854 -0.0025 1.0000 0.0295
-7.000 -0.7623 0.01480 0.00806 -0.0021 1.0000 0.0310
-6.750 -0.7372 0.01436 0.00759 -0.0017 1.0000 0.0330
-6.500 -0.7135 0.01371 0.00691 -0.0011 1.0000 0.0354
-6.250 -0.6882 0.01337 0.00660 -0.0007 1.0000 0.0385
-6.000 -0.6637 0.01293 0.00614 -0.0002 1.0000 0.0423
-5.750 -0.6390 0.01260 0.00585 0.0003 1.0000 0.0463
-5.500 -0.6147 0.01231 0.00554 0.0009 1.0000 0.0499
-5.250 -0.5924 0.01187 0.00515 0.0018 1.0000 0.0544
-5.000 -0.5701 0.01167 0.00494 0.0028 1.0000 0.0580
-4.750 -0.5416 0.01125 0.00455 0.0023 0.9977 0.0631
-4.500 -0.5039 0.01096 0.00427 0.0000 0.9939 0.0690
-4.250 -0.4679 0.01056 0.00392 -0.0020 0.9876 0.0772
-4.000 -0.4319 0.01023 0.00362 -0.0039 0.9805 0.0871
-3.750 -0.3974 0.00985 0.00334 -0.0055 0.9714 0.1034
-3.500 -0.3663 0.00942 0.00307 -0.0064 0.9591 0.1374
-3.250 -0.3391 0.00886 0.00282 -0.0065 0.9450 0.2081
-3.000 -0.3150 0.00821 0.00257 -0.0060 0.9308 0.3097
-2.750 -0.2932 0.00734 0.00230 -0.0052 0.9176 0.4644
-2.500 -0.2720 0.00667 0.00224 -0.0038 0.9059 0.6249
-2.250 -0.2457 0.00661 0.00226 -0.0033 0.8945 0.6746
-2.000 -0.2188 0.00660 0.00225 -0.0029 0.8855 0.7013
-1.750 -0.1913 0.00660 0.00224 -0.0026 0.8772 0.7208
-1.500 -0.1637 0.00660 0.00224 -0.0023 0.8702 0.7353
-1.250 -0.1359 0.00658 0.00224 -0.0021 0.8622 0.7488
-1.000 -0.1086 0.00660 0.00223 -0.0017 0.8545 0.7614
-0.750 -0.0815 0.00659 0.00224 -0.0013 0.8441 0.7769
-0.500 -0.0547 0.00660 0.00226 -0.0007 0.8343 0.7932
-0.250 -0.0269 0.00662 0.00226 -0.0005 0.8263 0.8044
0.000 0.0000 0.00660 0.00228 0.0000 0.8158 0.8158
0.250 0.0269 0.00662 0.00227 0.0005 0.8045 0.8264
0.500 0.0547 0.00660 0.00226 0.0008 0.7930 0.8344
0.750 0.0815 0.00659 0.00224 0.0013 0.7771 0.8441
1.000 0.1086 0.00660 0.00223 0.0017 0.7614 0.8545
1.250 0.1359 0.00658 0.00224 0.0021 0.7487 0.8623
1.500 0.1637 0.00660 0.00224 0.0023 0.7353 0.8702
1.750 0.1912 0.00660 0.00224 0.0026 0.7208 0.8772
2.000 0.2187 0.00660 0.00225 0.0029 0.7014 0.8855
2.250 0.2457 0.00661 0.00226 0.0033 0.6746 0.8945
2.500 0.2719 0.00667 0.00224 0.0039 0.6249 0.9059
2.750 0.2932 0.00734 0.00230 0.0052 0.4640 0.9176
3.000 0.3150 0.00821 0.00257 0.0060 0.3096 0.9308
3.250 0.3391 0.00886 0.00282 0.0065 0.2082 0.9451
3.500 0.3663 0.00942 0.00307 0.0064 0.1374 0.9591
3.750 0.3974 0.00985 0.00334 0.0055 0.1033 0.9715
4.000 0.4320 0.01023 0.00362 0.0039 0.0871 0.9806
4.250 0.4680 0.01056 0.00392 0.0020 0.0772 0.9876
4.500 0.5040 0.01096 0.00427 0.0000 0.0690 0.9939
4.750 0.5417 0.01125 0.00455 -0.0023 0.0631 0.9978
5.000 0.5700 0.01167 0.00494 -0.0028 0.0580 1.0000
5.250 0.5924 0.01187 0.00515 -0.0018 0.0544 1.0000
5.500 0.6147 0.01231 0.00553 -0.0009 0.0499 1.0000
5.750 0.6390 0.01260 0.00585 -0.0003 0.0463 1.0000
6.000 0.6637 0.01293 0.00614 0.0002 0.0423 1.0000
6.250 0.6882 0.01337 0.00660 0.0007 0.0386 1.0000
6.500 0.7135 0.01371 0.00691 0.0011 0.0354 1.0000
6.750 0.7372 0.01436 0.00758 0.0017 0.0329 1.0000
7.000 0.7623 0.01480 0.00806 0.0021 0.0310 1.0000
7.250 0.7870 0.01529 0.00854 0.0025 0.0295 1.0000
7.500 0.8090 0.01632 0.00958 0.0033 0.0281 1.0000
7.750 0.8336 0.01690 0.01025 0.0038 0.0273 1.0000
8.000 0.8576 0.01760 0.01102 0.0043 0.0265 1.0000
8.250 0.8815 0.01830 0.01178 0.0049 0.0256 1.0000
8.500 0.9054 0.01892 0.01244 0.0053 0.0248 1.0000
8.750 0.9285 0.01967 0.01321 0.0058 0.0241 1.0000
9.000 0.9482 0.02144 0.01507 0.0067 0.0233 1.0000
9.250 0.9706 0.02244 0.01622 0.0074 0.0230 1.0000
9.500 0.9921 0.02357 0.01752 0.0082 0.0227 1.0000
9.750 1.0125 0.02490 0.01903 0.0090 0.0223 1.0000
10.000 1.0314 0.02642 0.02076 0.0100 0.0220 1.0000
10.250 1.0486 0.02815 0.02271 0.0111 0.0217 1.0000
10.500 1.0635 0.03012 0.02493 0.0124 0.0214 1.0000
10.750 1.0752 0.03238 0.02747 0.0139 0.0212 1.0000
11.000 1.0828 0.03501 0.03040 0.0157 0.0210 1.0000
11.250 1.0884 0.03748 0.03313 0.0175 0.0207 1.0000
11.500 1.0944 0.03950 0.03532 0.0192 0.0205 1.0000
11.750 1.0987 0.04140 0.03736 0.0209 0.0202 1.0000
12.000 1.0929 0.04365 0.03978 0.0237 0.0201 1.0000
12.250 1.0809 0.04657 0.04290 0.0259 0.0200 1.0000
12.500 0.6810 0.09468 0.09239 0.0075 0.0240 1.0000
12.750 0.6841 0.09713 0.09482 0.0065 0.0237 1.0000
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Polar data table (+)
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