XFOIL Version 6.96 Calculated polar for: YS-900 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.6209 0.09498 0.09294 -0.0188 1.0000 0.0076 -9.000 -0.6397 0.08686 0.08485 -0.0253 1.0000 0.0060 -8.750 -0.6517 0.08466 0.08264 -0.0238 1.0000 0.0067 -8.500 -0.6731 0.08045 0.07838 -0.0225 1.0000 0.0062 -8.250 -0.6910 0.07901 0.07692 -0.0185 1.0000 0.0067 -8.000 -0.7123 0.07658 0.07445 -0.0136 1.0000 0.0065 -7.750 -0.7239 0.07400 0.07182 -0.0103 1.0000 0.0069 -7.500 -0.7342 0.07027 0.06800 -0.0067 1.0000 0.0065 -7.250 -0.7415 0.06783 0.06550 -0.0034 1.0000 0.0073 -7.000 -0.7475 0.06443 0.06198 0.0003 1.0000 0.0071 -6.750 -0.7508 0.06189 0.05934 0.0036 1.0000 0.0084 -6.500 -0.7524 0.05884 0.05617 0.0072 1.0000 0.0089 -6.250 -0.7528 0.05565 0.05283 0.0111 1.0000 0.0093 -6.000 -0.7511 0.05260 0.04962 0.0148 1.0000 0.0095 -5.750 -0.7479 0.04954 0.04639 0.0187 1.0000 0.0098 -5.500 -0.7424 0.04665 0.04330 0.0225 1.0000 0.0103 -5.250 -0.7353 0.04387 0.04031 0.0263 1.0000 0.0108 -5.000 -0.7042 0.04105 0.03720 0.0263 0.9966 0.0120 -4.750 -0.6679 0.04003 0.03589 0.0255 0.9929 0.0131 -4.500 -0.6483 0.03777 0.03333 0.0273 0.9892 0.0132 -4.250 -0.6271 0.03546 0.03066 0.0287 0.9862 0.0133 -4.000 -0.6177 0.02869 0.02338 0.0318 0.9842 0.0149 -3.750 -0.5915 0.02658 0.02109 0.0315 0.9828 0.0167 -3.500 -0.5717 0.02489 0.01920 0.0333 0.9782 0.0190 -3.250 -0.5454 0.02368 0.01773 0.0343 0.9756 0.0228 -1.750 -0.3661 0.01545 0.00846 0.0354 0.9650 0.0114 -1.500 -0.3425 0.01464 0.00753 0.0366 0.9619 0.0096 -1.250 -0.3124 0.01425 0.00701 0.0362 0.9597 0.0085 -1.000 -0.2790 0.01413 0.00675 0.0349 0.9582 0.0080 -0.750 -0.2435 0.01420 0.00670 0.0330 0.9570 0.0083 -0.500 -0.0781 0.01563 0.01169 0.0060 0.9791 0.9716 -0.250 -0.0375 0.01558 0.01161 0.0027 0.9774 0.9735 0.000 0.0021 0.01556 0.01159 -0.0005 0.9754 0.9757 0.250 0.0416 0.01559 0.01162 -0.0036 0.9732 0.9778 0.500 0.0819 0.01565 0.01171 -0.0068 0.9713 0.9795 0.750 0.2477 0.01414 0.00661 -0.0340 0.0088 0.9572 1.000 0.2829 0.01409 0.00669 -0.0358 0.0079 0.9584 1.250 0.3160 0.01418 0.00690 -0.0371 0.0083 0.9600 1.500 0.3455 0.01456 0.00742 -0.0374 0.0094 0.9623 1.750 0.3689 0.01527 0.00825 -0.0362 0.0108 0.9658 2.000 0.3847 0.01672 0.00979 -0.0332 0.0130 0.9704 4.500 0.6513 0.03759 0.03313 -0.0281 0.0133 0.9896 4.750 0.6748 0.04005 0.03591 -0.0272 0.0131 0.9946 5.000 0.7059 0.04115 0.03729 -0.0269 0.0123 0.9970 5.250 0.7360 0.04365 0.04007 -0.0268 0.0111 1.0000 5.500 0.7444 0.04627 0.04290 -0.0231 0.0105 1.0000 5.750 0.7496 0.04923 0.04606 -0.0192 0.0101 1.0000 6.000 0.7535 0.05219 0.04920 -0.0154 0.0095 1.0000 6.250 0.7555 0.05521 0.05238 -0.0116 0.0089 1.0000 6.500 0.7562 0.05773 0.05502 -0.0082 0.0074 1.0000 6.750 0.7545 0.06035 0.05774 -0.0047 0.0069 1.0000 7.000 0.7504 0.06335 0.06086 -0.0010 0.0066 1.0000 7.250 0.7432 0.06565 0.06321 0.0023 0.0061 1.0000 7.500 0.7364 0.06892 0.06659 0.0061 0.0061 1.0000 7.750 0.7239 0.07202 0.06976 0.0098 0.0058 1.0000 8.000 0.7068 0.07548 0.07329 0.0139 0.0057 1.0000 8.250 0.6913 0.07757 0.07543 0.0185 0.0057 1.0000 8.500 0.6694 0.07987 0.07777 0.0230 0.0057 1.0000 8.750 0.6571 0.08252 0.08046 0.0252 0.0059 1.0000 9.000 0.6419 0.08624 0.08422 0.0258 0.0060 1.0000 9.250 0.6283 0.09096 0.08896 0.0236 0.0060 1.0000