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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5680 0.10208 0.09491 0.0040 1.0000 0.2911 -8.750 -0.5704 0.09949 0.09239 0.0046 1.0000 0.3073 -8.500 -0.5796 0.09752 0.09054 0.0054 1.0000 0.3230 -8.250 -0.5551 0.09253 0.08550 0.0074 1.0000 0.3434 -8.000 -0.5651 0.09087 0.08397 0.0090 1.0000 0.3643 -7.750 -0.5451 0.08662 0.07969 0.0109 1.0000 0.3879 -7.500 -0.5408 0.08408 0.07720 0.0136 1.0000 0.4182 -7.250 -0.5458 0.08238 0.07561 0.0170 1.0000 0.4486 -6.750 -0.5198 0.05112 0.04300 -0.0500 1.0000 0.1560 -6.500 -0.4869 0.04437 0.03571 -0.0540 1.0000 0.1311 -6.250 -0.4519 0.03960 0.03023 -0.0575 1.0000 0.1244 -6.000 -0.4178 0.03563 0.02564 -0.0600 1.0000 0.1240 -5.750 -0.3837 0.03206 0.02147 -0.0616 1.0000 0.1218 -5.500 -0.3512 0.02917 0.01806 -0.0622 1.0000 0.1227 -5.250 -0.3214 0.02685 0.01540 -0.0623 1.0000 0.1317 -5.000 -0.2940 0.02492 0.01323 -0.0615 1.0000 0.1440 -4.750 -0.2699 0.02317 0.01146 -0.0596 1.0000 0.1579 -4.500 -0.2449 0.02148 0.00996 -0.0586 1.0000 0.1948 -4.250 -0.2195 0.01765 0.00910 -0.0570 1.0000 0.6103 -4.000 -0.2348 0.01873 0.01037 -0.0423 1.0000 0.7634 -3.750 -0.2447 0.01897 0.01055 -0.0307 1.0000 0.8283 -3.500 -0.2530 0.01848 0.00998 -0.0201 1.0000 0.8849 -3.250 -0.1729 0.01710 0.00783 -0.0256 1.0000 1.0000 -3.000 -0.1632 0.01663 0.00720 -0.0238 1.0000 1.0000 -2.750 -0.1516 0.01621 0.00663 -0.0222 1.0000 1.0000 -2.500 -0.1336 0.01590 0.00614 -0.0218 1.0000 1.0000 -2.250 -0.1095 0.01568 0.00574 -0.0225 1.0000 1.0000 -2.000 -0.0814 0.01555 0.00540 -0.0238 1.0000 1.0000 -1.750 -0.0510 0.01548 0.00518 -0.0255 1.0000 1.0000 -1.500 -0.0194 0.01546 0.00504 -0.0273 1.0000 1.0000 -1.250 0.0129 0.01547 0.00497 -0.0292 1.0000 1.0000 -1.000 0.0455 0.01553 0.00495 -0.0310 1.0000 1.0000 -0.750 0.0780 0.01560 0.00500 -0.0328 1.0000 1.0000 -0.500 0.1104 0.01571 0.00510 -0.0345 1.0000 1.0000 -0.250 0.1426 0.01584 0.00527 -0.0361 1.0000 1.0000 0.000 0.1745 0.01599 0.00550 -0.0377 1.0000 1.0000 0.250 0.2060 0.01616 0.00577 -0.0391 1.0000 1.0000 0.500 0.2372 0.01635 0.00610 -0.0404 1.0000 1.0000 0.750 0.2679 0.01657 0.00649 -0.0416 1.0000 1.0000 1.000 0.2983 0.01682 0.00698 -0.0427 1.0000 1.0000 1.250 0.3283 0.01709 0.00750 -0.0437 1.0000 1.0000 1.500 0.3579 0.01739 0.00810 -0.0446 1.0000 1.0000 1.750 0.3871 0.01772 0.00880 -0.0454 1.0000 1.0000 2.000 0.4160 0.01810 0.00963 -0.0462 1.0000 1.0000 2.250 0.4449 0.01851 0.01074 -0.0467 1.0000 1.0000 2.500 0.5760 0.02453 0.01309 -0.0578 0.1425 1.0000 2.750 0.6150 0.02704 0.01573 -0.0587 0.1310 1.0000 3.000 0.6481 0.02952 0.01847 -0.0591 0.1194 1.0000 3.250 0.6806 0.03217 0.02157 -0.0590 0.1153 1.0000 3.500 0.7120 0.03537 0.02533 -0.0586 0.1169 1.0000 3.750 0.7409 0.03905 0.02951 -0.0582 0.1209 1.0000 4.000 0.7675 0.04313 0.03390 -0.0578 0.1230 1.0000 4.250 0.7929 0.04696 0.03879 -0.0567 0.1334 1.0000 4.500 0.8191 0.05257 0.04527 -0.0567 0.1613 1.0000 5.750 0.8506 0.09100 0.08600 -0.1010 0.4286 1.0000 6.000 0.8739 0.09505 0.09007 -0.0959 0.3881 1.0000 6.250 0.8723 0.09872 0.09368 -0.0940 0.3593 1.0000 6.500 0.8812 0.10287 0.09781 -0.0915 0.3318 1.0000