XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0606 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.750 -0.6488 0.08096 0.07867 -0.0142 1.0000 0.0070 -9.500 -0.6594 0.07431 0.07207 -0.0186 1.0000 0.0069 -9.250 -0.6765 0.06145 0.05920 -0.0334 1.0000 0.0067 -9.000 -0.6822 0.05224 0.04979 -0.0400 1.0000 0.0067 -8.750 -0.6786 0.04439 0.04164 -0.0440 1.0000 0.0068 -8.500 -0.6677 0.03648 0.03330 -0.0473 1.0000 0.0070 -8.250 -0.6486 0.02938 0.02562 -0.0503 1.0000 0.0073 -8.000 -0.6245 0.02666 0.02259 -0.0515 1.0000 0.0077 -7.750 -0.5984 0.02384 0.01938 -0.0528 1.0000 0.0083 -7.500 -0.5698 0.02032 0.01535 -0.0544 1.0000 0.0085 -7.250 -0.5412 0.01791 0.01256 -0.0554 1.0000 0.0087 -7.000 -0.5130 0.01622 0.01060 -0.0560 1.0000 0.0091 -6.750 -0.4849 0.01499 0.00916 -0.0565 1.0000 0.0096 -6.500 -0.4571 0.01406 0.00808 -0.0568 1.0000 0.0100 -6.250 -0.4293 0.01333 0.00721 -0.0571 1.0000 0.0103 -6.000 -0.4000 0.01235 0.00610 -0.0579 1.0000 0.0107 -5.750 -0.3699 0.01145 0.00510 -0.0588 1.0000 0.0116 -5.500 -0.3416 0.01100 0.00462 -0.0592 1.0000 0.0128 -5.250 -0.3129 0.01057 0.00414 -0.0596 1.0000 0.0138 -5.000 -0.2840 0.01018 0.00366 -0.0601 1.0000 0.0149 -4.750 -0.2554 0.00986 0.00329 -0.0604 1.0000 0.0166 -4.500 -0.2262 0.00952 0.00295 -0.0609 1.0000 0.0224 -4.250 -0.1980 0.00929 0.00273 -0.0611 1.0000 0.0307 -4.000 -0.1699 0.00910 0.00257 -0.0614 1.0000 0.0408 -3.750 -0.1411 0.00887 0.00241 -0.0618 1.0000 0.0618 -3.500 -0.1104 0.00848 0.00225 -0.0629 1.0000 0.1163 -3.250 -0.0785 0.00805 0.00212 -0.0642 1.0000 0.1987 -3.000 -0.0456 0.00761 0.00200 -0.0658 1.0000 0.2987 -2.750 -0.0107 0.00708 0.00191 -0.0680 1.0000 0.4289 -2.500 0.0227 0.00672 0.00193 -0.0695 1.0000 0.5455 -2.250 0.0527 0.00659 0.00200 -0.0701 1.0000 0.6055 -2.000 0.0814 0.00656 0.00208 -0.0704 1.0000 0.6395 -1.750 0.1097 0.00656 0.00215 -0.0705 1.0000 0.6640 -1.500 0.1392 0.00656 0.00221 -0.0709 0.9993 0.6795 -1.250 0.1825 0.00638 0.00209 -0.0743 0.9929 0.6957 -1.000 0.2244 0.00620 0.00198 -0.0773 0.9848 0.7128 -0.750 0.2640 0.00602 0.00187 -0.0797 0.9723 0.7253 -0.500 0.3068 0.00581 0.00172 -0.0828 0.9480 0.7342 -0.250 0.3415 0.00637 0.00151 -0.0833 0.7318 0.7418 0.000 0.3615 0.00758 0.00176 -0.0817 0.4884 0.7493 0.250 0.3866 0.00830 0.00196 -0.0814 0.3434 0.7564 0.500 0.4125 0.00890 0.00215 -0.0813 0.2283 0.7630 0.750 0.4388 0.00946 0.00236 -0.0812 0.1323 0.7696 1.000 0.4652 0.00996 0.00258 -0.0810 0.0645 0.7761 1.250 0.4926 0.01025 0.00281 -0.0810 0.0421 0.7826 1.500 0.5198 0.01048 0.00304 -0.0808 0.0312 0.7881 1.750 0.5472 0.01075 0.00330 -0.0807 0.0226 0.7944 2.000 0.5741 0.01112 0.00368 -0.0804 0.0163 0.8007 2.250 0.6013 0.01145 0.00407 -0.0802 0.0148 0.8080 2.500 0.6280 0.01184 0.00458 -0.0798 0.0136 0.8140 2.750 0.6545 0.01230 0.00512 -0.0794 0.0128 0.8200 3.000 0.6810 0.01273 0.00559 -0.0791 0.0117 0.8257 3.250 0.7067 0.01338 0.00632 -0.0786 0.0107 0.8322 3.500 0.7312 0.01456 0.00766 -0.0778 0.0101 0.8384 3.750 0.7568 0.01524 0.00848 -0.0773 0.0098 0.8447 4.000 0.7825 0.01603 0.00941 -0.0768 0.0095 0.8513 4.250 0.8076 0.01701 0.01057 -0.0761 0.0092 0.8578 4.500 0.8328 0.01829 0.01206 -0.0754 0.0087 0.8641 4.750 0.8568 0.01993 0.01399 -0.0744 0.0084 0.8696 5.000 0.8800 0.02214 0.01656 -0.0733 0.0082 0.8758 5.250 0.9013 0.02487 0.01974 -0.0718 0.0080 0.8818 5.500 0.9227 0.02696 0.02219 -0.0707 0.0075 0.8888 5.750 0.9436 0.02870 0.02420 -0.0696 0.0071 0.8958 6.000 0.9601 0.03223 0.02817 -0.0678 0.0068 0.9034 6.250 0.9726 0.03682 0.03323 -0.0656 0.0066 0.9116 6.500 0.9812 0.04237 0.03923 -0.0634 0.0064 0.9212 6.750 0.9819 0.04978 0.04710 -0.0609 0.0062 0.9344 7.000 0.9740 0.05880 0.05654 -0.0588 0.0061 0.9673 7.250 0.9654 0.06845 0.06646 -0.0592 0.0060 1.0000 7.500 0.9604 0.07598 0.07416 -0.0608 0.0060 1.0000 7.750 0.9517 0.08325 0.08152 -0.0637 0.0060 1.0000 8.000 0.9409 0.08936 0.08770 -0.0667 0.0060 1.0000