XFOIL Version 6.96 Calculated polar for: S1010 HPV 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6687 0.08700 0.08228 0.0061 1.0000 0.1256 -8.000 -0.6973 0.08197 0.07723 -0.0040 1.0000 0.1290 -7.750 -0.6879 0.07678 0.07209 -0.0030 1.0000 0.1319 -7.500 -0.6779 0.07297 0.06827 -0.0023 1.0000 0.1357 -7.250 -0.7024 0.05355 0.04801 -0.0142 1.0000 0.0739 -7.000 -0.7051 0.04355 0.03692 -0.0143 1.0000 0.0635 -6.750 -0.6908 0.03908 0.03188 -0.0132 1.0000 0.0629 -6.500 -0.6766 0.03409 0.02642 -0.0122 1.0000 0.0640 -6.250 -0.6556 0.03261 0.02498 -0.0116 1.0000 0.0708 -6.000 -0.6357 0.02916 0.02080 -0.0102 1.0000 0.0740 -5.750 -0.6141 0.02604 0.01721 -0.0091 1.0000 0.0781 -5.500 -0.5907 0.02468 0.01556 -0.0081 1.0000 0.0880 -5.250 -0.5671 0.02284 0.01363 -0.0072 1.0000 0.0982 -5.000 -0.5431 0.02133 0.01196 -0.0063 1.0000 0.1118 -4.750 -0.5191 0.02026 0.01068 -0.0054 1.0000 0.1292 -4.500 -0.4955 0.01911 0.00960 -0.0046 1.0000 0.1471 -4.250 -0.4712 0.01807 0.00856 -0.0038 1.0000 0.1645 -4.000 -0.4472 0.01720 0.00764 -0.0029 1.0000 0.1861 -3.750 -0.4236 0.01621 0.00681 -0.0020 1.0000 0.2076 -3.500 -0.4004 0.01536 0.00609 -0.0010 1.0000 0.2358 -3.250 -0.3779 0.01449 0.00540 0.0000 1.0000 0.2718 -3.000 -0.3566 0.01355 0.00481 0.0013 1.0000 0.3261 -2.750 -0.3388 0.01221 0.00431 0.0033 1.0000 0.4525 -2.500 -0.3225 0.01093 0.00465 0.0085 1.0000 0.8351 -2.250 -0.2235 0.01150 0.00484 -0.0025 1.0000 0.9702 -2.000 -0.1389 0.01131 0.00429 -0.0139 1.0000 1.0000 -1.750 -0.1199 0.01114 0.00399 -0.0126 1.0000 1.0000 -1.500 -0.1012 0.01099 0.00376 -0.0112 1.0000 1.0000 -1.250 -0.0830 0.01086 0.00357 -0.0097 1.0000 1.0000 -1.000 -0.0653 0.01076 0.00343 -0.0080 1.0000 1.0000 -0.750 -0.0483 0.01068 0.00331 -0.0061 1.0000 1.0000 -0.500 -0.0320 0.01062 0.00323 -0.0041 1.0000 1.0000 -0.250 -0.0160 0.01059 0.00318 -0.0021 1.0000 1.0000 0.000 0.0000 0.01058 0.00316 0.0000 1.0000 1.0000 0.250 0.0160 0.01059 0.00318 0.0021 1.0000 1.0000 0.500 0.0320 0.01062 0.00323 0.0041 1.0000 1.0000 0.750 0.0483 0.01068 0.00331 0.0061 1.0000 1.0000 1.000 0.0654 0.01076 0.00342 0.0080 1.0000 1.0000 1.250 0.0830 0.01086 0.00357 0.0096 1.0000 1.0000 1.500 0.1013 0.01099 0.00376 0.0112 1.0000 1.0000 1.750 0.1199 0.01113 0.00399 0.0126 1.0000 1.0000 2.000 0.1389 0.01131 0.00429 0.0138 1.0000 1.0000 2.250 0.2233 0.01150 0.00484 0.0025 0.9703 1.0000 2.500 0.3225 0.01092 0.00465 -0.0085 0.8351 1.0000 2.750 0.3389 0.01221 0.00431 -0.0033 0.4526 1.0000 3.000 0.3566 0.01354 0.00481 -0.0013 0.3262 1.0000 3.250 0.3780 0.01449 0.00540 -0.0001 0.2719 1.0000 3.500 0.4005 0.01536 0.00609 0.0010 0.2358 1.0000 3.750 0.4236 0.01622 0.00681 0.0020 0.2075 1.0000 4.000 0.4473 0.01719 0.00764 0.0029 0.1861 1.0000 4.250 0.4713 0.01807 0.00856 0.0038 0.1645 1.0000 4.500 0.4955 0.01912 0.00961 0.0046 0.1472 1.0000 4.750 0.5191 0.02026 0.01068 0.0054 0.1293 1.0000 5.000 0.5431 0.02133 0.01197 0.0063 0.1119 1.0000 5.250 0.5672 0.02284 0.01363 0.0072 0.0982 1.0000 5.500 0.5907 0.02469 0.01557 0.0081 0.0880 1.0000 5.750 0.6141 0.02605 0.01722 0.0091 0.0782 1.0000 6.000 0.6358 0.02916 0.02081 0.0102 0.0741 1.0000 6.250 0.6557 0.03261 0.02497 0.0116 0.0708 1.0000 6.500 0.6767 0.03410 0.02642 0.0122 0.0639 1.0000 6.750 0.6909 0.03908 0.03188 0.0131 0.0629 1.0000 7.000 0.7053 0.04351 0.03688 0.0142 0.0635 1.0000 7.250 0.7033 0.05422 0.04875 0.0142 0.0782 1.0000 7.500 0.7163 0.05901 0.05336 0.0146 0.0758 1.0000 7.750 0.6876 0.07681 0.07212 0.0027 0.1320 1.0000 8.000 0.7005 0.08193 0.07717 0.0047 0.1292 1.0000 8.250 0.6697 0.08700 0.08228 -0.0060 0.1257 1.0000