XFOIL Version 6.96 Calculated polar for: HQ 1.5/8 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.4984 0.08750 0.08408 -0.0205 1.0000 0.0385 -8.250 -0.5024 0.08296 0.07961 -0.0251 1.0000 0.0395 -8.000 -0.5096 0.07806 0.07477 -0.0310 1.0000 0.0399 -7.750 -0.5128 0.07305 0.06968 -0.0367 1.0000 0.0402 -7.500 -0.5138 0.06907 0.06553 -0.0399 1.0000 0.0406 -7.250 -0.5117 0.06543 0.06170 -0.0413 1.0000 0.0408 -7.000 -0.5158 0.05797 0.05418 -0.0426 1.0000 0.0417 -6.750 -0.5069 0.05403 0.05032 -0.0421 1.0000 0.0428 -6.500 -0.4977 0.05101 0.04728 -0.0416 1.0000 0.0442 -6.250 -0.4882 0.04792 0.04410 -0.0411 1.0000 0.0461 -6.000 -0.4778 0.04468 0.04065 -0.0406 1.0000 0.0489 -5.750 -0.4696 0.04145 0.03671 -0.0395 1.0000 0.0546 -5.500 -0.4583 0.03772 0.03307 -0.0384 1.0000 0.0564 -5.250 -0.4459 0.03526 0.03054 -0.0369 1.0000 0.0587 -5.000 -0.4242 0.02822 0.02225 -0.0329 1.0000 0.0293 -4.750 -0.4117 0.02600 0.02010 -0.0322 1.0000 0.0350 -4.500 -0.3906 0.02234 0.01566 -0.0298 1.0000 0.0303 -4.250 -0.3698 0.02122 0.01422 -0.0281 1.0000 0.0291 -4.000 -0.3490 0.01968 0.01245 -0.0267 1.0000 0.0288 -3.750 -0.3268 0.01726 0.00977 -0.0255 1.0000 0.0294 -3.500 -0.3051 0.01545 0.00789 -0.0244 1.0000 0.0316 -3.250 -0.2690 0.01466 0.00705 -0.0262 0.9962 0.0395 -3.000 -0.2334 0.01325 0.00566 -0.0279 0.9922 0.0566 -2.750 -0.1979 0.01208 0.00481 -0.0300 0.9871 0.1276 -2.500 -0.1707 0.00997 0.00470 -0.0311 0.9829 0.5923 -2.250 -0.1389 0.00982 0.00475 -0.0316 0.9761 0.6920 -2.000 -0.1033 0.00979 0.00476 -0.0328 0.9705 0.7481 -1.750 -0.0707 0.00975 0.00474 -0.0334 0.9639 0.7909 -1.500 -0.0365 0.00967 0.00466 -0.0343 0.9579 0.8264 -1.250 -0.0032 0.00954 0.00458 -0.0347 0.9523 0.8675 -1.000 0.0317 0.00937 0.00447 -0.0353 0.9458 0.9072 -0.750 0.0817 0.00926 0.00434 -0.0393 0.9430 0.9392 -0.500 0.1356 0.00916 0.00416 -0.0445 0.9399 0.9590 -0.250 0.1851 0.00903 0.00398 -0.0490 0.9331 0.9740 0.000 0.2364 0.00885 0.00376 -0.0538 0.9281 0.9860 0.250 0.2797 0.00871 0.00359 -0.0572 0.9184 1.0000 0.500 0.3029 0.00860 0.00342 -0.0564 0.9044 1.0000 0.750 0.3255 0.00857 0.00334 -0.0554 0.8910 1.0000 1.000 0.3489 0.00859 0.00330 -0.0544 0.8783 1.0000 1.250 0.3730 0.00861 0.00331 -0.0535 0.8659 1.0000 1.500 0.3974 0.00865 0.00332 -0.0526 0.8530 1.0000 1.750 0.4221 0.00869 0.00333 -0.0517 0.8399 1.0000 2.000 0.4468 0.00872 0.00335 -0.0507 0.8256 1.0000 2.250 0.4715 0.00876 0.00338 -0.0498 0.8100 1.0000 2.500 0.4966 0.00879 0.00344 -0.0488 0.7940 1.0000 2.750 0.5216 0.00883 0.00348 -0.0479 0.7762 1.0000 3.000 0.5463 0.00888 0.00353 -0.0469 0.7542 1.0000 3.250 0.5709 0.00893 0.00356 -0.0458 0.7279 1.0000 3.500 0.5951 0.00901 0.00358 -0.0446 0.6953 1.0000 3.750 0.6188 0.00915 0.00368 -0.0434 0.6525 1.0000 4.000 0.6417 0.00939 0.00376 -0.0421 0.5940 1.0000 4.250 0.6628 0.00985 0.00390 -0.0405 0.5118 1.0000 4.500 0.6830 0.01053 0.00420 -0.0391 0.4270 1.0000 4.750 0.7040 0.01125 0.00461 -0.0380 0.3548 1.0000 5.000 0.7246 0.01207 0.00515 -0.0370 0.2966 1.0000 5.250 0.7473 0.01270 0.00562 -0.0363 0.2372 1.0000 5.500 0.7672 0.01375 0.00625 -0.0353 0.1463 1.0000 5.750 0.7853 0.01520 0.00727 -0.0340 0.0895 1.0000 6.000 0.8062 0.01628 0.00833 -0.0329 0.0679 1.0000 6.250 0.8270 0.01736 0.00937 -0.0320 0.0498 1.0000 6.500 0.8480 0.01845 0.01047 -0.0310 0.0321 1.0000 6.750 0.8658 0.02045 0.01256 -0.0291 0.0250 1.0000 7.000 0.8866 0.02206 0.01431 -0.0278 0.0216 1.0000 7.250 0.9070 0.02434 0.01670 -0.0266 0.0200 1.0000 7.500 0.9274 0.02776 0.02040 -0.0254 0.0192 1.0000 7.750 0.9470 0.03093 0.02393 -0.0240 0.0191 1.0000 8.000 0.9647 0.03395 0.02733 -0.0225 0.0193 1.0000 8.250 0.9782 0.03746 0.03128 -0.0207 0.0193 1.0000 8.500 0.9847 0.04193 0.03620 -0.0187 0.0191 1.0000 8.750 0.9818 0.04752 0.04230 -0.0164 0.0188 1.0000 10.250 0.8964 0.07438 0.07103 -0.0088 0.0239 1.0000 10.500 0.8732 0.08173 0.07849 -0.0138 0.0251 1.0000