XFOIL Version 6.96 Calculated polar for: ONERA HOR04 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 ------ -------- --------- --------- -------- -------- -------- -8.500 -0.6526 0.10659 0.10503 0.0326 1.0000 0.0059 -8.250 -0.6486 0.10296 0.10142 0.0312 1.0000 0.0060 -8.000 -0.6447 0.09939 0.09786 0.0296 1.0000 0.0061 -7.750 -0.6415 0.09582 0.09430 0.0277 1.0000 0.0062 -7.500 -0.6352 0.09198 0.09047 0.0246 1.0000 0.0063 -7.250 -0.6247 0.08773 0.08622 0.0202 1.0000 0.0064 -7.000 -0.6116 0.08323 0.08172 0.0151 1.0000 0.0066 -6.750 -0.5958 0.07848 0.07696 0.0094 1.0000 0.0068 -6.500 -0.5771 0.07347 0.07192 0.0032 1.0000 0.0071 -6.250 -0.5554 0.06822 0.06662 -0.0033 1.0000 0.0074 -6.000 -0.5304 0.06271 0.06105 -0.0098 1.0000 0.0081 -5.750 -0.4942 0.05730 0.05552 -0.0166 1.0000 0.0093 -5.500 -0.4645 0.05203 0.05012 -0.0212 1.0000 0.0094 -5.250 -0.4349 0.04678 0.04471 -0.0250 1.0000 0.0095 -5.000 -0.4055 0.04203 0.03976 -0.0280 1.0000 0.0095 -4.750 -0.3761 0.03339 0.03075 -0.0331 1.0000 0.0103 -4.500 -0.3491 0.03113 0.02837 -0.0346 1.0000 0.0108 -4.250 -0.3197 0.02846 0.02553 -0.0360 1.0000 0.0113 -4.000 -0.2889 0.02574 0.02260 -0.0373 1.0000 0.0121 -3.750 -0.2574 0.02319 0.01983 -0.0382 1.0000 0.0132 -3.500 -0.2267 0.02248 0.01899 -0.0380 1.0000 0.0151 -3.250 -0.1956 0.02095 0.01724 -0.0384 1.0000 0.0155 -3.000 -0.1563 0.01320 0.00864 -0.0406 1.0000 0.0117 -2.750 -0.1260 0.01195 0.00724 -0.0410 1.0000 0.0129 -2.500 -0.0959 0.01071 0.00586 -0.0413 1.0000 0.0132 -2.250 -0.0664 0.00977 0.00483 -0.0415 1.0000 0.0136 -2.000 -0.0370 0.00904 0.00406 -0.0418 1.0000 0.0141 -1.750 -0.0072 0.00815 0.00308 -0.0422 1.0000 0.0142 -1.500 0.0224 0.00758 0.00247 -0.0425 1.0000 0.0149 -1.250 0.0521 0.00702 0.00184 -0.0428 1.0000 0.0203 -1.000 0.0808 0.00689 0.00174 -0.0430 1.0000 0.0250 -0.750 0.1097 0.00665 0.00150 -0.0432 1.0000 0.0332 -0.500 0.1384 0.00647 0.00137 -0.0434 1.0000 0.0512 -0.250 0.1676 0.00516 0.00131 -0.0446 1.0000 0.4855 0.000 0.2008 0.00451 0.00149 -0.0461 0.9812 0.7520 0.250 0.2192 0.00393 0.00140 -0.0432 0.9211 1.0000 0.500 0.2403 0.00427 0.00136 -0.0412 0.8270 1.0000 0.750 0.2667 0.00457 0.00136 -0.0407 0.7572 1.0000 1.000 0.2941 0.00490 0.00138 -0.0406 0.6810 1.0000 1.250 0.3220 0.00528 0.00143 -0.0407 0.6016 1.0000 1.500 0.3502 0.00552 0.00151 -0.0408 0.5582 1.0000 1.750 0.3783 0.00583 0.00159 -0.0410 0.4869 1.0000 2.000 0.4064 0.00658 0.00175 -0.0415 0.3322 1.0000 2.250 0.4346 0.00751 0.00202 -0.0421 0.1591 1.0000 2.500 0.4629 0.00815 0.00227 -0.0425 0.0641 1.0000 2.750 0.4913 0.00847 0.00250 -0.0427 0.0452 1.0000 3.000 0.5195 0.00866 0.00272 -0.0428 0.0407 1.0000 3.250 0.5479 0.00900 0.00305 -0.0429 0.0344 1.0000 3.500 0.5761 0.00932 0.00341 -0.0430 0.0315 1.0000 3.750 0.6041 0.00944 0.00352 -0.0431 0.0289 1.0000 4.000 0.6322 0.00969 0.00378 -0.0432 0.0257 1.0000 4.250 0.6602 0.01042 0.00462 -0.0433 0.0212 1.0000 4.500 0.6881 0.01033 0.00449 -0.0434 0.0198 1.0000 4.750 0.7159 0.01046 0.00459 -0.0434 0.0163 1.0000 5.000 0.7437 0.01111 0.00530 -0.0435 0.0125 1.0000 5.250 0.7715 0.01147 0.00570 -0.0435 0.0108 1.0000 5.500 0.7990 0.01195 0.00622 -0.0435 0.0091 1.0000 5.750 0.8255 0.01391 0.00847 -0.0434 0.0079 1.0000 6.000 0.8526 0.01448 0.00913 -0.0433 0.0074 1.0000 6.250 0.8792 0.01561 0.01042 -0.0431 0.0071 1.0000 6.500 0.9054 0.01718 0.01222 -0.0429 0.0068 1.0000 6.750 0.9309 0.01907 0.01438 -0.0426 0.0065 1.0000 7.000 0.9560 0.02093 0.01654 -0.0424 0.0061 1.0000 7.250 0.9808 0.02242 0.01825 -0.0422 0.0057 1.0000 7.500 1.0053 0.02362 0.01962 -0.0421 0.0053 1.0000 7.750 1.0186 0.03122 0.02808 -0.0411 0.0049 1.0000 11.000 0.7459 0.13257 0.13114 -0.0631 0.0080 1.0000 11.250 0.7432 0.13694 0.13551 -0.0644 0.0080 1.0000