XFOIL Version 6.96 Calculated polar for: ONERA OA213 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.750 -0.4170 0.12463 0.11896 -0.0071 1.0000 0.1933 -9.500 -0.3769 0.11740 0.11169 -0.0041 1.0000 0.2041 -9.250 -0.4039 0.11715 0.11155 -0.0107 1.0000 0.2101 -9.000 -0.3666 0.11072 0.10508 -0.0080 1.0000 0.2230 -8.750 -0.3594 0.10693 0.10135 -0.0095 1.0000 0.2320 -8.500 -0.3764 0.10569 0.10020 -0.0136 1.0000 0.2432 -8.250 -0.3536 0.10107 0.09560 -0.0124 1.0000 0.2579 -8.000 -0.3415 0.09743 0.09199 -0.0123 1.0000 0.2717 -7.750 -0.3372 0.09451 0.08913 -0.0123 1.0000 0.2873 -7.500 -0.3373 0.09210 0.08679 -0.0114 1.0000 0.3048 -7.250 -0.3829 0.09277 0.08763 -0.0090 1.0000 0.3100 -7.000 -0.3524 0.08874 0.08362 -0.0059 1.0000 0.3391 -6.750 -0.3654 0.08750 0.08246 -0.0020 1.0000 0.3567 -6.500 -0.4140 0.08805 0.08319 0.0029 1.0000 0.3599 -6.250 -0.3760 0.08431 0.07944 0.0066 1.0000 0.4063 -6.000 -0.4024 0.08386 0.07912 0.0122 1.0000 0.4262 -5.750 -0.4099 0.08262 0.07797 0.0178 1.0000 0.4609 -5.500 -0.3278 0.07762 0.07288 0.0202 1.0000 0.5576 -4.500 -0.2369 0.06715 0.06248 0.0274 1.0000 0.7464 -4.250 -0.3005 0.06727 0.06279 0.0342 1.0000 0.6980 -4.000 -0.3602 0.06722 0.06291 0.0417 1.0000 0.6757 -3.750 -0.4125 0.04960 0.04256 -0.0132 1.0000 0.1745 -3.500 -0.3844 0.04562 0.03790 -0.0139 1.0000 0.1467 -3.250 -0.3605 0.04317 0.03504 -0.0138 1.0000 0.1375 -3.000 -0.3367 0.04100 0.03241 -0.0137 1.0000 0.1343 -2.750 -0.3128 0.03921 0.03021 -0.0136 1.0000 0.1333 -2.500 -0.2839 0.03753 0.02808 -0.0140 0.9986 0.1321 -2.250 -0.2336 0.03608 0.02622 -0.0181 0.9895 0.1375 -2.000 -0.1836 0.03493 0.02466 -0.0219 0.9797 0.1459 -1.750 -0.1353 0.03408 0.02374 -0.0253 0.9690 0.1633 -1.500 -0.0880 0.03350 0.02313 -0.0283 0.9576 0.1936 -1.250 0.0069 0.02984 0.02261 -0.0347 0.9482 1.0000 -1.000 0.0451 0.03076 0.02293 -0.0373 0.9327 1.0000 -0.750 0.0860 0.03171 0.02348 -0.0405 0.9160 1.0000 -0.500 0.1340 0.03273 0.02414 -0.0446 0.8986 1.0000 -0.250 0.1621 0.03349 0.02467 -0.0453 0.8802 1.0000 0.000 0.1944 0.03424 0.02521 -0.0465 0.8613 1.0000 0.250 0.2347 0.03494 0.02572 -0.0487 0.8422 1.0000 0.500 0.2852 0.03553 0.02613 -0.0523 0.8232 1.0000 0.750 0.3230 0.03597 0.02642 -0.0535 0.8042 1.0000 1.000 0.3419 0.03640 0.02675 -0.0519 0.7838 1.0000 1.250 0.3705 0.03666 0.02691 -0.0514 0.7639 1.0000 1.500 0.4039 0.03668 0.02684 -0.0511 0.7447 1.0000 1.750 0.4378 0.03644 0.02653 -0.0504 0.7265 1.0000 2.000 0.4704 0.03598 0.02598 -0.0491 0.7090 1.0000 2.250 0.4856 0.03614 0.02609 -0.0465 0.6864 1.0000 2.500 0.5130 0.03561 0.02550 -0.0444 0.6675 1.0000 2.750 0.5417 0.03477 0.02460 -0.0421 0.6497 1.0000 3.000 0.5705 0.03380 0.02356 -0.0396 0.6325 1.0000 3.250 0.5991 0.03276 0.02244 -0.0369 0.6154 1.0000 3.500 0.6274 0.03166 0.02126 -0.0342 0.5983 1.0000 3.750 0.6501 0.03133 0.02085 -0.0319 0.5771 1.0000 4.000 0.6747 0.03087 0.02029 -0.0296 0.5570 1.0000 4.250 0.7008 0.03031 0.01962 -0.0274 0.5381 1.0000 4.500 0.7268 0.02996 0.01912 -0.0254 0.5199 1.0000 4.750 0.7524 0.02982 0.01880 -0.0237 0.5022 1.0000 5.000 0.7757 0.03021 0.01909 -0.0224 0.4838 1.0000 5.250 0.7985 0.03086 0.01969 -0.0215 0.4666 1.0000 5.500 0.8223 0.03145 0.02018 -0.0206 0.4512 1.0000 5.750 0.8474 0.03195 0.02054 -0.0196 0.4373 1.0000 6.000 0.8691 0.03303 0.02164 -0.0191 0.4234 1.0000 6.250 0.8887 0.03453 0.02325 -0.0189 0.4109 1.0000 6.500 0.9095 0.03589 0.02464 -0.0184 0.4000 1.0000 6.750 0.9373 0.03641 0.02501 -0.0177 0.3904 1.0000 7.000 0.9497 0.03879 0.02766 -0.0177 0.3800 1.0000 7.250 0.9734 0.04005 0.02889 -0.0173 0.3726 1.0000 7.500 0.9795 0.04318 0.03233 -0.0172 0.3647 1.0000 7.750 1.0119 0.04361 0.03260 -0.0167 0.3582 1.0000 8.000 0.9916 0.04933 0.03882 -0.0167 0.3517 1.0000 8.250 1.0052 0.05156 0.04115 -0.0163 0.3456 1.0000 8.500 1.0151 0.05446 0.04414 -0.0161 0.3411 1.0000 8.750 0.7430 0.09506 0.08484 -0.0324 0.3559 1.0000 9.000 0.7520 0.09953 0.08937 -0.0337 0.3586 1.0000