XFOIL Version 6.96 Calculated polar for: USA 48 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 ------ -------- --------- --------- -------- -------- -------- -18.500 -0.7633 0.12495 0.12272 -0.0426 1.0000 0.0107 -18.250 -0.8031 0.11319 0.11079 -0.0485 1.0000 0.0106 -18.000 -0.8340 0.10392 0.10139 -0.0532 1.0000 0.0106 -17.750 -0.8540 0.09672 0.09408 -0.0568 1.0000 0.0106 -17.500 -0.8756 0.08955 0.08680 -0.0604 1.0000 0.0106 -17.250 -0.8981 0.08250 0.07963 -0.0639 1.0000 0.0106 -17.000 -0.9134 0.07673 0.07376 -0.0667 1.0000 0.0107 -16.750 -0.9340 0.07048 0.06740 -0.0697 1.0000 0.0107 -16.500 -0.9450 0.06571 0.06254 -0.0719 1.0000 0.0107 -16.250 -0.9638 0.06009 0.05679 -0.0744 1.0000 0.0107 -16.000 -0.9801 0.05489 0.05148 -0.0767 1.0000 0.0107 -15.750 -0.9890 0.05076 0.04725 -0.0783 1.0000 0.0107 -15.500 -1.0137 0.04489 0.04122 -0.0806 1.0000 0.0109 -15.250 -1.0268 0.04096 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0.7351 0.00909 0.00303 -0.0523 0.2933 0.7401 4.750 0.7535 0.00938 0.00323 -0.0505 0.2691 0.7458 5.000 0.7724 0.00964 0.00342 -0.0488 0.2494 0.7494 5.250 0.7883 0.01003 0.00367 -0.0466 0.2198 0.7570 5.500 0.7965 0.01074 0.00410 -0.0430 0.1599 0.7611 5.750 0.8110 0.01117 0.00443 -0.0406 0.1368 0.7648 6.000 0.8310 0.01137 0.00463 -0.0392 0.1332 0.7686 6.250 0.8502 0.01162 0.00487 -0.0376 0.1297 0.7725 6.500 0.8690 0.01185 0.00512 -0.0360 0.1259 0.7782 6.750 0.8878 0.01206 0.00537 -0.0344 0.1227 0.7849 7.000 0.9074 0.01224 0.00559 -0.0329 0.1217 0.7926 7.250 0.9267 0.01238 0.00582 -0.0315 0.1203 0.8030 7.500 0.9443 0.01252 0.00607 -0.0296 0.1187 0.8220 7.750 0.9880 0.01261 0.00650 -0.0336 0.1161 0.9281 8.000 1.0742 0.01322 0.00713 -0.0472 0.1113 0.9744 8.250 1.0929 0.01363 0.00755 -0.0457 0.1084 1.0000 8.500 1.1139 0.01383 0.00777 -0.0447 0.1070 1.0000 8.750 1.1349 0.01403 0.00800 -0.0437 0.1052 1.0000 9.000 1.1542 0.01431 0.00829 -0.0424 0.1016 1.0000 9.250 1.1715 0.01470 0.00865 -0.0408 0.0969 1.0000 9.500 1.1919 0.01494 0.00890 -0.0398 0.0905 1.0000 10.000 1.2004 0.01704 0.01066 -0.0331 0.0358 1.0000 10.250 1.2111 0.01782 0.01144 -0.0308 0.0307 1.0000 10.500 1.2247 0.01847 0.01211 -0.0290 0.0278 1.0000 10.750 1.2356 0.01929 0.01294 -0.0270 0.0252 1.0000 11.000 1.2490 0.02001 0.01368 -0.0253 0.0232 1.0000 11.250 1.2611 0.02081 0.01451 -0.0236 0.0216 1.0000 11.500 1.2694 0.02189 0.01562 -0.0216 0.0198 1.0000 11.750 1.2829 0.02267 0.01645 -0.0203 0.0190 1.0000 12.000 1.2945 0.02359 0.01741 -0.0188 0.0181 1.0000 12.250 1.3048 0.02465 0.01850 -0.0173 0.0174 1.0000 12.500 1.3112 0.02602 0.01991 -0.0156 0.0164 1.0000 12.750 1.3182 0.02741 0.02137 -0.0141 0.0158 1.0000 13.000 1.3281 0.02862 0.02264 -0.0130 0.0152 1.0000 13.250 1.3368 0.02998 0.02405 -0.0119 0.0147 1.0000 13.500 1.3461 0.03133 0.02545 -0.0110 0.0141 1.0000 13.750 1.3522 0.03300 0.02717 -0.0100 0.0137 1.0000 14.000 1.3552 0.03500 0.02923 -0.0091 0.0133 1.0000 14.250 1.3517 0.03770 0.03202 -0.0080 0.0127 1.0000 14.500 1.3551 0.03982 0.03423 -0.0074 0.0126 1.0000 14.750 1.3587 0.04198 0.03647 -0.0068 0.0123 1.0000 15.000 1.3637 0.04406 0.03861 -0.0065 0.0120 1.0000 15.250 1.3653 0.04655 0.04119 -0.0062 0.0118 1.0000 15.500 1.3664 0.04916 0.04387 -0.0060 0.0115 1.0000 15.750 1.3669 0.05187 0.04665 -0.0060 0.0112 1.0000 16.000 1.3672 0.05468 0.04953 -0.0060 0.0110 1.0000 16.250 1.3651 0.05786 0.05278 -0.0062 0.0108 1.0000 16.500 1.3602 0.06143 0.05645 -0.0066 0.0106 1.0000 16.750 1.3527 0.06540 0.06049 -0.0071 0.0103 1.0000 17.000 1.3435 0.06969 0.06487 -0.0079 0.0101 1.0000 17.250 1.3232 0.07554 0.07086 -0.0090 0.0100 1.0000 17.500 1.3173 0.07955 0.07496 -0.0099 0.0099 1.0000 17.750 1.3167 0.08291 0.07840 -0.0107 0.0098 1.0000 18.000 1.3077 0.08749 0.08307 -0.0119 0.0098 1.0000 18.250 1.3039 0.09142 0.08709 -0.0130 0.0096 1.0000 18.500 1.2972 0.09577 0.09153 -0.0143 0.0095 1.0000 18.750 1.2913 0.10005 0.09590 -0.0156 0.0093 1.0000 19.000 1.2825 0.10481 0.10075 -0.0172 0.0093 1.0000 19.250 1.2743 0.10955 0.10557 -0.0188 0.0092 1.0000