NASA SC(2)-0610 AIRFOIL (sc20610-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0610 AIRFOIL (sc20610-il) Reynolds number: 100,000 Max Cl/Cd: 31.82 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20610-il-100000-n5.txt Download as CSV file: xf-sc20610-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0610 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6866 0.08095 0.07508 -0.0433 1.0000 0.0345
-11.500 -0.7120 0.07277 0.06681 -0.0489 1.0000 0.0343
-11.250 -0.7373 0.06644 0.06036 -0.0523 1.0000 0.0341
-11.000 -0.7620 0.06162 0.05541 -0.0533 1.0000 0.0340
-10.750 -0.7861 0.05765 0.05129 -0.0528 1.0000 0.0340
-10.500 -0.8036 0.05317 0.04650 -0.0532 1.0000 0.0342
-10.250 -0.8147 0.04861 0.04144 -0.0532 1.0000 0.0347
-10.000 -0.8065 0.04634 0.03911 -0.0528 1.0000 0.0353
-9.750 -0.7960 0.04428 0.03691 -0.0524 1.0000 0.0361
-9.500 -0.7845 0.04191 0.03431 -0.0521 1.0000 0.0369
-9.250 -0.7714 0.03918 0.03126 -0.0519 1.0000 0.0377
-9.000 -0.7556 0.03652 0.02824 -0.0517 1.0000 0.0386
-8.750 -0.7375 0.03404 0.02538 -0.0515 1.0000 0.0397
-8.500 -0.7175 0.03177 0.02271 -0.0512 1.0000 0.0411
-8.250 -0.6965 0.03009 0.02084 -0.0509 1.0000 0.0429
-8.000 -0.6749 0.02903 0.01974 -0.0507 1.0000 0.0450
-7.750 -0.6524 0.02780 0.01835 -0.0504 1.0000 0.0475
-7.500 -0.6295 0.02646 0.01677 -0.0498 1.0000 0.0499
-7.250 -0.6069 0.02524 0.01551 -0.0494 1.0000 0.0525
-7.000 -0.5835 0.02441 0.01468 -0.0492 1.0000 0.0562
-6.750 -0.5596 0.02353 0.01365 -0.0488 1.0000 0.0603
-6.500 -0.5357 0.02257 0.01274 -0.0486 1.0000 0.0644
-6.250 -0.5110 0.02186 0.01199 -0.0485 1.0000 0.0699
-6.000 -0.4857 0.02105 0.01117 -0.0485 1.0000 0.0754
-5.750 -0.4598 0.02041 0.01054 -0.0487 1.0000 0.0826
-5.500 -0.4330 0.01972 0.00987 -0.0490 1.0000 0.0901
-5.250 -0.4059 0.01919 0.00930 -0.0494 1.0000 0.1006
-5.000 -0.3775 0.01855 0.00874 -0.0501 1.0000 0.1135
-4.750 -0.3482 0.01793 0.00827 -0.0511 1.0000 0.1340
-4.500 -0.3174 0.01726 0.00783 -0.0525 1.0000 0.1753
-4.250 -0.2807 0.01614 0.00736 -0.0558 1.0000 0.2907
-4.000 -0.2493 0.01527 0.00769 -0.0571 1.0000 0.5067
-3.750 -0.2309 0.01566 0.00832 -0.0543 1.0000 0.5902
-3.500 -0.2058 0.01596 0.00857 -0.0535 1.0000 0.6304
-3.250 -0.1811 0.01629 0.00884 -0.0527 1.0000 0.6580
-3.000 -0.1616 0.01677 0.00931 -0.0504 1.0000 0.6779
-2.750 -0.1432 0.01728 0.00980 -0.0478 1.0000 0.6962
-2.500 -0.1260 0.01781 0.01033 -0.0449 1.0000 0.7135
-2.250 -0.1090 0.01827 0.01080 -0.0421 1.0000 0.7288
-2.000 -0.0893 0.01863 0.01115 -0.0401 1.0000 0.7424
-1.750 -0.0677 0.01886 0.01136 -0.0388 1.0000 0.7521
-1.500 -0.0438 0.01899 0.01147 -0.0381 1.0000 0.7588
-1.250 -0.0202 0.01910 0.01158 -0.0375 1.0000 0.7643
-1.000 0.0151 0.01919 0.01162 -0.0397 0.9976 0.7707
-0.750 0.0489 0.01929 0.01174 -0.0410 0.9940 0.7746
-0.500 0.0821 0.01934 0.01181 -0.0424 0.9892 0.7781
-0.250 0.1210 0.01941 0.01190 -0.0450 0.9846 0.7816
0.000 0.1597 0.01943 0.01195 -0.0477 0.9786 0.7853
0.250 0.2028 0.01942 0.01199 -0.0511 0.9721 0.7881
0.500 0.2406 0.01930 0.01195 -0.0531 0.9616 0.7905
0.750 0.2854 0.01900 0.01175 -0.0561 0.9473 0.7929
1.000 0.3392 0.01836 0.01121 -0.0603 0.9271 0.7952
1.250 0.3868 0.01761 0.01057 -0.0630 0.9011 0.7977
1.500 0.4276 0.01699 0.01003 -0.0645 0.8674 0.8006
1.750 0.4614 0.01657 0.00968 -0.0647 0.8213 0.8033
2.000 0.5113 0.01607 0.00879 -0.0668 0.6712 0.8050
2.250 0.5332 0.01734 0.00875 -0.0649 0.4180 0.8071
2.500 0.5528 0.01857 0.00915 -0.0637 0.2596 0.8099
3.000 0.6036 0.02023 0.01009 -0.0636 0.1339 0.8161
3.250 0.6288 0.02088 0.01062 -0.0634 0.1140 0.8188
3.500 0.6527 0.02144 0.01118 -0.0626 0.1008 0.8214
3.750 0.6773 0.02205 0.01180 -0.0621 0.0908 0.8242
4.000 0.7023 0.02280 0.01250 -0.0618 0.0828 0.8269
4.250 0.7293 0.02346 0.01325 -0.0618 0.0755 0.8297
4.500 0.7560 0.02438 0.01409 -0.0620 0.0694 0.8325
4.750 0.7809 0.02511 0.01497 -0.0614 0.0641 0.8350
5.000 0.8047 0.02591 0.01575 -0.0608 0.0595 0.8378
5.250 0.8303 0.02698 0.01694 -0.0604 0.0558 0.8408
5.500 0.8568 0.02791 0.01799 -0.0603 0.0517 0.8438
5.750 0.8831 0.02904 0.01906 -0.0606 0.0486 0.8467
6.000 0.9097 0.03050 0.02079 -0.0603 0.0464 0.8493
6.250 0.9341 0.03195 0.02249 -0.0596 0.0441 0.8519
6.500 0.9579 0.03316 0.02384 -0.0592 0.0419 0.8547
6.750 0.9813 0.03441 0.02513 -0.0589 0.0401 0.8579
7.000 1.0042 0.03659 0.02775 -0.0583 0.0385 0.8612
7.250 1.0241 0.03909 0.03071 -0.0572 0.0373 0.8644
7.500 1.0404 0.04167 0.03373 -0.0556 0.0364 0.8676
7.750 1.0548 0.04448 0.03697 -0.0540 0.0356 0.8711
8.000 1.0672 0.04747 0.04036 -0.0524 0.0350 0.8746
8.250 1.0798 0.05010 0.04329 -0.0512 0.0343 0.8780
8.500 1.0899 0.05225 0.04565 -0.0495 0.0336 0.8812
8.750 1.0995 0.05451 0.04804 -0.0481 0.0330 0.8850
9.000 1.0885 0.05954 0.05370 -0.0449 0.0325 0.8888
9.250 1.0722 0.06473 0.05939 -0.0421 0.0321 0.8927
9.500 1.0502 0.06922 0.06423 -0.0389 0.0321 0.8967
9.750 1.0255 0.07415 0.06942 -0.0370 0.0321 0.9013
10.000 1.0000 0.08002 0.07553 -0.0376 0.0322 0.9053
10.250 0.9723 0.08735 0.08306 -0.0410 0.0324 0.9089
10.500 0.9469 0.09680 0.09265 -0.0479 0.0326 0.9123
10.750 0.9258 0.10907 0.10498 -0.0570 0.0329 0.9152
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0610 AIRFOIL (sc20610-il)