NASA SC(2)-0010 AIRFOIL (sc20010-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0010 AIRFOIL (sc20010-il) Reynolds number: 1,000,000 Max Cl/Cd: 67.71 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20010-il-1000000.txt Download as CSV file: xf-sc20010-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.750 -1.1585 0.08015 0.07779 -0.0061 1.0000 0.0133
-15.500 -1.2086 0.06522 0.06254 -0.0169 1.0000 0.0132
-15.250 -1.2335 0.05728 0.05438 -0.0218 1.0000 0.0131
-15.000 -1.2499 0.05167 0.04860 -0.0245 1.0000 0.0132
-14.750 -1.2623 0.04713 0.04389 -0.0259 1.0000 0.0132
-14.500 -1.2707 0.04343 0.04005 -0.0265 1.0000 0.0133
-14.250 -1.2764 0.04034 0.03681 -0.0264 1.0000 0.0133
-14.000 -1.2791 0.03779 0.03413 -0.0257 1.0000 0.0134
-13.750 -1.2808 0.03554 0.03174 -0.0244 1.0000 0.0135
-13.500 -1.2804 0.03364 0.02971 -0.0227 1.0000 0.0136
-13.250 -1.2786 0.03199 0.02794 -0.0204 1.0000 0.0137
-13.000 -1.2754 0.03057 0.02639 -0.0179 1.0000 0.0138
-12.750 -1.2674 0.02922 0.02492 -0.0158 1.0000 0.0140
-12.500 -1.2558 0.02791 0.02347 -0.0143 1.0000 0.0141
-12.250 -1.2422 0.02668 0.02212 -0.0128 1.0000 0.0143
-12.000 -1.2270 0.02552 0.02084 -0.0116 1.0000 0.0145
-11.750 -1.2099 0.02448 0.01969 -0.0105 1.0000 0.0147
-11.500 -1.1913 0.02357 0.01867 -0.0095 1.0000 0.0149
-11.250 -1.1716 0.02273 0.01772 -0.0086 1.0000 0.0151
-11.000 -1.1514 0.02189 0.01678 -0.0078 1.0000 0.0153
-10.750 -1.1305 0.02108 0.01587 -0.0071 1.0000 0.0154
-10.500 -1.1088 0.02034 0.01505 -0.0064 1.0000 0.0155
-10.250 -1.0922 0.01872 0.01328 -0.0051 1.0000 0.0158
-10.000 -1.0721 0.01766 0.01213 -0.0042 1.0000 0.0162
-9.750 -1.0499 0.01689 0.01131 -0.0035 1.0000 0.0165
-9.500 -1.0265 0.01627 0.01065 -0.0030 1.0000 0.0168
-9.250 -1.0026 0.01570 0.01003 -0.0026 1.0000 0.0172
-9.000 -0.9783 0.01517 0.00945 -0.0021 1.0000 0.0175
-8.750 -0.9538 0.01466 0.00890 -0.0017 1.0000 0.0180
-8.500 -0.9289 0.01417 0.00837 -0.0014 1.0000 0.0184
-8.250 -0.9037 0.01374 0.00788 -0.0010 1.0000 0.0188
-8.000 -0.8782 0.01334 0.00744 -0.0007 1.0000 0.0192
-7.750 -0.8543 0.01265 0.00673 -0.0002 1.0000 0.0203
-7.500 -0.8288 0.01224 0.00632 0.0001 1.0000 0.0211
-7.250 -0.8029 0.01189 0.00594 0.0003 1.0000 0.0220
-7.000 -0.7768 0.01157 0.00560 0.0006 1.0000 0.0230
-6.750 -0.7515 0.01112 0.00515 0.0009 1.0000 0.0249
-6.500 -0.7255 0.01082 0.00486 0.0012 1.0000 0.0271
-6.250 -0.6999 0.01049 0.00455 0.0016 1.0000 0.0303
-6.000 -0.6742 0.01028 0.00436 0.0019 1.0000 0.0336
-5.750 -0.6493 0.00999 0.00412 0.0024 0.9999 0.0376
-5.500 -0.6142 0.00982 0.00394 0.0007 0.9954 0.0407
-5.250 -0.5805 0.00951 0.00368 -0.0008 0.9881 0.0455
-5.000 -0.5469 0.00935 0.00351 -0.0021 0.9781 0.0486
-4.750 -0.5162 0.00910 0.00327 -0.0028 0.9634 0.0528
-4.500 -0.4892 0.00894 0.00310 -0.0025 0.9440 0.0566
-4.250 -0.4639 0.00885 0.00295 -0.0018 0.9237 0.0590
-4.000 -0.4385 0.00865 0.00274 -0.0012 0.9047 0.0655
-3.750 -0.4120 0.00854 0.00258 -0.0008 0.8879 0.0712
-3.250 -0.3585 0.00811 0.00225 -0.0004 0.8599 0.1081
-3.000 -0.3314 0.00787 0.00210 -0.0003 0.8485 0.1393
-2.750 -0.3048 0.00754 0.00194 -0.0002 0.8387 0.1908
-2.500 -0.2777 0.00715 0.00178 -0.0002 0.8305 0.2580
-2.250 -0.2515 0.00666 0.00162 -0.0001 0.8225 0.3558
-2.000 -0.2257 0.00599 0.00144 -0.0001 0.8142 0.4907
-1.750 -0.2001 0.00549 0.00135 0.0002 0.8050 0.6178
-1.500 -0.1719 0.00538 0.00135 0.0003 0.7956 0.6617
-1.250 -0.1435 0.00538 0.00134 0.0003 0.7886 0.6838
-1.000 -0.1145 0.00534 0.00134 0.0001 0.7834 0.6982
-0.750 -0.0857 0.00534 0.00132 0.0000 0.7757 0.7085
-0.500 -0.0570 0.00535 0.00132 0.0000 0.7655 0.7208
-0.250 -0.0286 0.00534 0.00132 0.0000 0.7572 0.7346
0.000 0.0000 0.00531 0.00134 0.0000 0.7493 0.7491
0.250 0.0285 0.00534 0.00132 0.0000 0.7346 0.7572
0.500 0.0570 0.00535 0.00132 0.0000 0.7211 0.7653
0.750 0.0857 0.00534 0.00133 0.0000 0.7087 0.7758
1.000 0.1145 0.00534 0.00134 -0.0001 0.6983 0.7834
1.250 0.1435 0.00538 0.00134 -0.0002 0.6837 0.7887
1.500 0.1719 0.00539 0.00135 -0.0002 0.6614 0.7957
1.750 0.2001 0.00549 0.00135 -0.0002 0.6176 0.8052
2.000 0.2257 0.00599 0.00144 0.0001 0.4901 0.8143
2.250 0.2515 0.00666 0.00162 0.0001 0.3558 0.8225
2.500 0.2777 0.00715 0.00178 0.0002 0.2581 0.8305
2.750 0.3048 0.00754 0.00194 0.0002 0.1907 0.8387
3.000 0.3314 0.00787 0.00210 0.0003 0.1394 0.8485
3.250 0.3585 0.00811 0.00225 0.0004 0.1081 0.8600
3.750 0.4120 0.00854 0.00258 0.0008 0.0712 0.8879
4.000 0.4385 0.00865 0.00274 0.0012 0.0655 0.9048
4.250 0.4639 0.00885 0.00295 0.0018 0.0590 0.9238
4.500 0.4892 0.00894 0.00310 0.0025 0.0566 0.9441
4.750 0.5162 0.00910 0.00327 0.0028 0.0528 0.9635
5.000 0.5469 0.00935 0.00351 0.0021 0.0486 0.9781
5.250 0.5805 0.00951 0.00368 0.0008 0.0454 0.9881
5.500 0.6142 0.00982 0.00394 -0.0007 0.0407 0.9954
5.750 0.6494 0.00999 0.00412 -0.0024 0.0375 0.9999
6.000 0.6742 0.01028 0.00436 -0.0019 0.0336 1.0000
6.250 0.6999 0.01049 0.00455 -0.0016 0.0304 1.0000
6.500 0.7255 0.01083 0.00487 -0.0012 0.0272 1.0000
6.750 0.7515 0.01112 0.00515 -0.0009 0.0249 1.0000
7.000 0.7768 0.01157 0.00560 -0.0006 0.0230 1.0000
7.250 0.8029 0.01189 0.00594 -0.0003 0.0221 1.0000
7.500 0.8288 0.01224 0.00632 -0.0001 0.0211 1.0000
7.750 0.8544 0.01265 0.00673 0.0002 0.0203 1.0000
8.000 0.8782 0.01333 0.00744 0.0007 0.0192 1.0000
8.250 0.9037 0.01374 0.00789 0.0010 0.0188 1.0000
8.500 0.9290 0.01417 0.00837 0.0014 0.0184 1.0000
8.750 0.9538 0.01465 0.00890 0.0017 0.0180 1.0000
9.000 0.9784 0.01516 0.00945 0.0021 0.0175 1.0000
9.250 1.0027 0.01570 0.01003 0.0026 0.0171 1.0000
9.500 1.0266 0.01627 0.01064 0.0030 0.0168 1.0000
9.750 1.0499 0.01689 0.01131 0.0035 0.0165 1.0000
10.000 1.0721 0.01766 0.01214 0.0042 0.0162 1.0000
10.250 1.0922 0.01872 0.01328 0.0051 0.0158 1.0000
10.500 1.1088 0.02035 0.01506 0.0064 0.0155 1.0000
10.750 1.1306 0.02108 0.01588 0.0071 0.0154 1.0000
11.000 1.1515 0.02189 0.01678 0.0078 0.0153 1.0000
11.250 1.1717 0.02272 0.01771 0.0086 0.0151 1.0000
11.500 1.1915 0.02356 0.01866 0.0095 0.0149 1.0000
11.750 1.2101 0.02448 0.01968 0.0104 0.0147 1.0000
12.000 1.2270 0.02553 0.02085 0.0116 0.0145 1.0000
12.250 1.2424 0.02668 0.02212 0.0128 0.0143 1.0000
12.500 1.2560 0.02791 0.02347 0.0142 0.0141 1.0000
12.750 1.2677 0.02921 0.02491 0.0158 0.0140 1.0000
13.000 1.2758 0.03056 0.02638 0.0178 0.0138 1.0000
13.250 1.2791 0.03199 0.02793 0.0204 0.0137 1.0000
13.500 1.2810 0.03362 0.02970 0.0226 0.0136 1.0000
13.750 1.2817 0.03550 0.03170 0.0243 0.0135 1.0000
14.000 1.2804 0.03772 0.03406 0.0255 0.0134 1.0000
14.250 1.2774 0.04030 0.03677 0.0262 0.0133 1.0000
14.500 1.2722 0.04334 0.03996 0.0264 0.0133 1.0000
14.750 1.2636 0.04707 0.04384 0.0258 0.0132 1.0000
15.000 1.2520 0.05151 0.04843 0.0244 0.0132 1.0000
15.250 1.2347 0.05726 0.05436 0.0217 0.0131 1.0000
15.500 1.2103 0.06514 0.06245 0.0167 0.0132 1.0000
15.750 1.1617 0.07975 0.07738 0.0062 0.0133 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0010 AIRFOIL (sc20010-il)