NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il) Reynolds number: 1,000,000 Max Cl/Cd: 96.57 at α=-0.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20706-il-1000000.txt Download as CSV file: xf-sc20706-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0706 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5030 0.08858 0.08700 -0.0175 1.0000 0.0103
-10.500 -0.5090 0.08370 0.08213 -0.0184 1.0000 0.0104
-10.250 -0.5150 0.07897 0.07741 -0.0193 1.0000 0.0105
-10.000 -0.5213 0.07438 0.07285 -0.0201 1.0000 0.0106
-9.750 -0.5295 0.06954 0.06802 -0.0211 1.0000 0.0106
-9.500 -0.5405 0.06454 0.06305 -0.0220 1.0000 0.0106
-9.250 -0.5570 0.05919 0.05772 -0.0231 1.0000 0.0106
-9.000 -0.5861 0.05251 0.05109 -0.0258 1.0000 0.0104
-8.750 -0.6005 0.04536 0.04384 -0.0328 1.0000 0.0103
-8.500 -0.6085 0.03857 0.03690 -0.0365 1.0000 0.0104
-8.250 -0.6076 0.03224 0.03036 -0.0397 1.0000 0.0106
-8.000 -0.5968 0.02703 0.02493 -0.0424 1.0000 0.0108
-7.750 -0.5903 0.03040 0.02759 -0.0511 1.0000 0.0092
-7.500 -0.5512 0.02173 0.01808 -0.0572 1.0000 0.0094
-7.250 -0.5205 0.01955 0.01557 -0.0587 1.0000 0.0097
-6.750 -0.4531 0.01457 0.00988 -0.0630 1.0000 0.0101
-6.500 -0.4158 0.01215 0.00720 -0.0658 1.0000 0.0109
-6.250 -0.3835 0.01130 0.00627 -0.0671 1.0000 0.0117
-6.000 -0.3527 0.01083 0.00574 -0.0680 1.0000 0.0126
-5.750 -0.3208 0.01032 0.00518 -0.0692 1.0000 0.0136
-5.500 -0.2891 0.00990 0.00469 -0.0703 1.0000 0.0144
-5.250 -0.2582 0.00958 0.00434 -0.0711 1.0000 0.0151
-5.000 -0.2212 0.00888 0.00352 -0.0734 1.0000 0.0186
-4.750 -0.1910 0.00868 0.00332 -0.0741 1.0000 0.0223
-4.500 -0.1591 0.00840 0.00301 -0.0752 1.0000 0.0286
-4.250 -0.1293 0.00826 0.00287 -0.0757 1.0000 0.0343
-4.000 -0.0971 0.00799 0.00270 -0.0769 1.0000 0.0608
-3.750 -0.0601 0.00749 0.00255 -0.0795 1.0000 0.1565
-3.500 -0.0167 0.00673 0.00240 -0.0837 1.0000 0.3327
-3.250 0.0249 0.00614 0.00232 -0.0874 1.0000 0.4949
-3.000 0.0579 0.00597 0.00235 -0.0887 1.0000 0.5672
-2.750 0.0877 0.00594 0.00239 -0.0892 1.0000 0.5990
-2.500 0.1170 0.00592 0.00245 -0.0895 1.0000 0.6248
-2.250 0.1460 0.00593 0.00254 -0.0898 1.0000 0.6530
-2.000 0.1742 0.00597 0.00261 -0.0899 1.0000 0.6681
-1.750 0.2023 0.00601 0.00271 -0.0900 1.0000 0.6850
-1.500 0.2479 0.00579 0.00254 -0.0939 0.9963 0.6974
-1.250 0.2996 0.00535 0.00217 -0.0991 0.9882 0.7112
-1.000 0.3480 0.00492 0.00180 -0.1035 0.9784 0.7207
-0.750 0.3901 0.00465 0.00157 -0.1064 0.9671 0.7296
-0.500 0.4307 0.00446 0.00139 -0.1090 0.9297 0.7364
-0.250 0.4462 0.00618 0.00158 -0.1059 0.5507 0.7429
0.000 0.4683 0.00731 0.00184 -0.1051 0.3159 0.7492
0.250 0.4926 0.00822 0.00208 -0.1047 0.1386 0.7562
0.500 0.5188 0.00873 0.00230 -0.1044 0.0586 0.7620
0.750 0.5460 0.00905 0.00252 -0.1043 0.0345 0.7684
1.000 0.5735 0.00926 0.00277 -0.1041 0.0280 0.7739
1.250 0.6004 0.00970 0.00324 -0.1038 0.0204 0.7797
1.500 0.6280 0.00989 0.00346 -0.1036 0.0184 0.7856
1.750 0.6551 0.01019 0.00379 -0.1034 0.0160 0.7911
2.000 0.6799 0.01124 0.00499 -0.1025 0.0133 0.7969
2.250 0.7068 0.01155 0.00535 -0.1022 0.0129 0.8023
2.500 0.7330 0.01209 0.00597 -0.1017 0.0125 0.8084
2.750 0.7591 0.01272 0.00667 -0.1012 0.0120 0.8142
3.000 0.7849 0.01339 0.00744 -0.1006 0.0114 0.8195
3.250 0.8111 0.01409 0.00824 -0.1001 0.0107 0.8252
3.500 0.8368 0.01484 0.00909 -0.0995 0.0100 0.8301
3.750 0.8621 0.01597 0.01035 -0.0988 0.0096 0.8354
4.000 0.8870 0.01729 0.01183 -0.0980 0.0092 0.8406
4.250 0.9091 0.01996 0.01483 -0.0966 0.0088 0.8451
4.500 0.9312 0.02249 0.01771 -0.0952 0.0088 0.8503
4.750 0.9519 0.02528 0.02089 -0.0936 0.0088 0.8552
5.000 0.9700 0.02872 0.02475 -0.0914 0.0089 0.8601
11.000 0.9387 0.17193 0.17071 -0.1224 0.0075 1.0000
11.250 0.9432 0.17649 0.17526 -0.1245 0.0069 1.0000
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Polar data table (+)
Polar graphs
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