NASA SC(2)-0010 AIRFOIL (sc20010-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA SC(2)-0010 AIRFOIL (sc20010-il) Reynolds number: 100,000 Max Cl/Cd: 28.66 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20010-il-100000-n5.txt Download as CSV file: xf-sc20010-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7733 0.09151 0.08604 -0.0078 1.0000 0.0339
-11.250 -0.7989 0.08087 0.07533 -0.0159 1.0000 0.0336
-11.000 -0.8260 0.07280 0.06713 -0.0214 1.0000 0.0334
-10.750 -0.8518 0.06665 0.06079 -0.0240 1.0000 0.0332
-10.500 -0.8757 0.06177 0.05570 -0.0239 1.0000 0.0332
-10.250 -0.8955 0.05736 0.05100 -0.0222 1.0000 0.0332
-10.000 -0.9085 0.05274 0.04598 -0.0204 1.0000 0.0334
-9.750 -0.9151 0.04829 0.04103 -0.0186 1.0000 0.0336
-9.500 -0.9132 0.04459 0.03688 -0.0169 1.0000 0.0341
-9.250 -0.8992 0.04284 0.03504 -0.0160 1.0000 0.0349
-9.000 -0.8845 0.04111 0.03315 -0.0151 1.0000 0.0360
-8.750 -0.8699 0.03900 0.03077 -0.0141 1.0000 0.0373
-8.500 -0.8550 0.03646 0.02783 -0.0129 1.0000 0.0386
-8.250 -0.8380 0.03386 0.02478 -0.0117 1.0000 0.0398
-8.000 -0.8184 0.03159 0.02211 -0.0107 1.0000 0.0411
-7.750 -0.7971 0.03010 0.02057 -0.0101 1.0000 0.0424
-7.500 -0.7751 0.02889 0.01926 -0.0095 1.0000 0.0445
-7.250 -0.7525 0.02760 0.01774 -0.0089 1.0000 0.0476
-7.000 -0.7296 0.02623 0.01623 -0.0082 1.0000 0.0503
-6.750 -0.7067 0.02515 0.01515 -0.0077 1.0000 0.0531
-6.500 -0.6832 0.02418 0.01403 -0.0070 1.0000 0.0573
-6.250 -0.6604 0.02313 0.01296 -0.0064 1.0000 0.0614
-6.000 -0.6374 0.02231 0.01213 -0.0057 1.0000 0.0663
-5.750 -0.6144 0.02144 0.01119 -0.0049 1.0000 0.0717
-5.500 -0.5921 0.02068 0.01049 -0.0042 1.0000 0.0775
-5.250 -0.5695 0.01997 0.00972 -0.0033 1.0000 0.0839
-5.000 -0.5480 0.01927 0.00907 -0.0024 1.0000 0.0908
-4.750 -0.5264 0.01863 0.00845 -0.0014 1.0000 0.0987
-4.500 -0.5052 0.01805 0.00790 -0.0004 1.0000 0.1083
-4.250 -0.4849 0.01746 0.00739 0.0008 1.0000 0.1203
-4.000 -0.4650 0.01690 0.00691 0.0020 1.0000 0.1365
-3.750 -0.4457 0.01632 0.00652 0.0033 1.0000 0.1631
-3.500 -0.4275 0.01570 0.00617 0.0046 1.0000 0.2114
-3.250 -0.4107 0.01489 0.00586 0.0060 1.0000 0.2995
-3.000 -0.3934 0.01377 0.00561 0.0073 0.9984 0.4687
-2.750 -0.3656 0.01333 0.00595 0.0079 0.9919 0.6520
-2.500 -0.3331 0.01334 0.00602 0.0072 0.9850 0.7129
-2.250 -0.3017 0.01339 0.00607 0.0068 0.9781 0.7546
-2.000 -0.2698 0.01348 0.00621 0.0067 0.9723 0.7902
-1.750 -0.2392 0.01360 0.00635 0.0068 0.9658 0.8204
-1.500 -0.2082 0.01377 0.00655 0.0071 0.9594 0.8484
-1.250 -0.1760 0.01394 0.00671 0.0071 0.9540 0.8738
-1.000 -0.1442 0.01405 0.00680 0.0067 0.9472 0.8907
-0.750 -0.1069 0.01412 0.00683 0.0050 0.9421 0.9015
-0.500 -0.0712 0.01412 0.00681 0.0034 0.9359 0.9095
-0.250 -0.0357 0.01414 0.00681 0.0018 0.9294 0.9168
0.000 0.0000 0.01413 0.00679 0.0000 0.9241 0.9241
0.250 0.0357 0.01414 0.00681 -0.0018 0.9168 0.9294
0.500 0.0712 0.01412 0.00681 -0.0034 0.9095 0.9359
0.750 0.1069 0.01412 0.00683 -0.0050 0.9015 0.9421
1.000 0.1442 0.01405 0.00680 -0.0067 0.8907 0.9472
1.250 0.1760 0.01394 0.00671 -0.0071 0.8738 0.9540
1.500 0.2083 0.01377 0.00655 -0.0071 0.8485 0.9594
1.750 0.2392 0.01360 0.00635 -0.0068 0.8205 0.9658
2.000 0.2698 0.01348 0.00621 -0.0067 0.7902 0.9723
2.250 0.3017 0.01339 0.00607 -0.0068 0.7545 0.9782
2.500 0.3331 0.01334 0.00601 -0.0072 0.7129 0.9851
2.750 0.3657 0.01333 0.00595 -0.0079 0.6520 0.9919
3.000 0.3934 0.01377 0.00561 -0.0073 0.4686 0.9985
3.250 0.4107 0.01489 0.00585 -0.0060 0.2996 1.0000
3.500 0.4274 0.01569 0.00617 -0.0046 0.2115 1.0000
3.750 0.4457 0.01632 0.00652 -0.0033 0.1631 1.0000
4.000 0.4649 0.01690 0.00691 -0.0020 0.1365 1.0000
4.250 0.4849 0.01746 0.00739 -0.0008 0.1203 1.0000
4.500 0.5052 0.01805 0.00789 0.0004 0.1083 1.0000
4.750 0.5263 0.01863 0.00845 0.0014 0.0987 1.0000
5.000 0.5479 0.01926 0.00907 0.0024 0.0908 1.0000
5.250 0.5695 0.01997 0.00972 0.0033 0.0839 1.0000
5.500 0.5921 0.02068 0.01048 0.0042 0.0775 1.0000
5.750 0.6144 0.02144 0.01119 0.0049 0.0717 1.0000
6.000 0.6374 0.02231 0.01213 0.0057 0.0663 1.0000
6.250 0.6604 0.02313 0.01297 0.0064 0.0614 1.0000
6.500 0.6833 0.02418 0.01403 0.0070 0.0573 1.0000
6.750 0.7068 0.02515 0.01515 0.0076 0.0531 1.0000
7.000 0.7296 0.02623 0.01623 0.0082 0.0503 1.0000
7.250 0.7525 0.02760 0.01774 0.0089 0.0476 1.0000
7.500 0.7752 0.02889 0.01925 0.0095 0.0445 1.0000
7.750 0.7971 0.03010 0.02057 0.0101 0.0424 1.0000
8.000 0.8184 0.03159 0.02211 0.0107 0.0411 1.0000
8.250 0.8380 0.03386 0.02478 0.0117 0.0398 1.0000
8.500 0.8550 0.03646 0.02783 0.0129 0.0386 1.0000
8.750 0.8700 0.03900 0.03077 0.0140 0.0373 1.0000
9.000 0.8846 0.04111 0.03315 0.0151 0.0360 1.0000
9.250 0.8993 0.04283 0.03504 0.0160 0.0349 1.0000
9.500 0.9134 0.04459 0.03688 0.0168 0.0341 1.0000
9.750 0.9152 0.04831 0.04105 0.0185 0.0336 1.0000
10.000 0.9087 0.05276 0.04599 0.0204 0.0334 1.0000
10.250 0.8957 0.05738 0.05102 0.0221 0.0332 1.0000
10.500 0.8759 0.06181 0.05574 0.0238 0.0332 1.0000
10.750 0.8521 0.06670 0.06085 0.0239 0.0332 1.0000
11.000 0.8262 0.07289 0.06722 0.0213 0.0334 1.0000
11.250 0.7993 0.08097 0.07545 0.0157 0.0336 1.0000
11.500 0.7736 0.09173 0.08626 0.0075 0.0339 1.0000
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Polar data table (+)
Polar graphs
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