XFOIL Version 6.96 Calculated polar for: ONERA NACA CAMBRE 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.4881 0.11312 0.10594 0.0151 1.0000 0.3538 -9.000 -0.4860 0.11044 0.10331 0.0157 1.0000 0.3723 -8.750 -0.4978 0.10886 0.10182 0.0164 1.0000 0.3907 -8.500 -0.4650 0.10377 0.09669 0.0175 1.0000 0.4118 -8.250 -0.4583 0.10115 0.09411 0.0188 1.0000 0.4356 -7.750 -0.4389 0.09461 0.08765 0.0206 1.0000 0.4752 -7.500 -0.4323 0.09148 0.08456 0.0212 1.0000 0.4909 -6.250 -0.5972 0.05035 0.04200 -0.0169 1.0000 0.1695 -6.000 -0.5871 0.04736 0.03818 -0.0148 1.0000 0.1555 -5.750 -0.5685 0.04367 0.03450 -0.0138 1.0000 0.1518 -5.500 -0.5519 0.04070 0.03114 -0.0121 1.0000 0.1456 -5.250 -0.5334 0.03879 0.02848 -0.0099 1.0000 0.1389 -5.000 -0.5124 0.03622 0.02575 -0.0086 1.0000 0.1364 -4.750 -0.4906 0.03408 0.02333 -0.0072 1.0000 0.1346 -4.500 -0.4684 0.03231 0.02128 -0.0058 1.0000 0.1353 -4.250 -0.4455 0.03079 0.01948 -0.0044 1.0000 0.1374 -4.000 -0.4213 0.02940 0.01784 -0.0032 1.0000 0.1391 -3.750 -0.3950 0.02784 0.01622 -0.0023 1.0000 0.1410 -3.500 -0.3684 0.02639 0.01489 -0.0015 1.0000 0.1460 -3.250 -0.3433 0.02542 0.01385 -0.0003 1.0000 0.1545 -3.000 -0.3193 0.02422 0.01279 0.0007 1.0000 0.1643 -2.750 -0.0218 0.02262 0.01367 -0.0277 1.0000 1.0000 -2.500 -0.0119 0.02201 0.01294 -0.0267 1.0000 1.0000 -2.250 -0.0043 0.02148 0.01234 -0.0251 1.0000 1.0000 -2.000 0.0003 0.02104 0.01185 -0.0230 1.0000 1.0000 -1.750 0.0011 0.02070 0.01148 -0.0202 1.0000 1.0000 -1.500 -0.0017 0.02045 0.01122 -0.0167 1.0000 1.0000 -1.250 -0.0086 0.02032 0.01106 -0.0125 1.0000 1.0000 -1.000 -0.0182 0.02028 0.01100 -0.0079 1.0000 1.0000 -0.750 -0.0280 0.02034 0.01101 -0.0032 1.0000 1.0000 -0.500 -0.0353 0.02051 0.01111 0.0010 1.0000 1.0000 -0.250 -0.0389 0.02078 0.01130 0.0045 1.0000 1.0000 0.000 -0.0388 0.02116 0.01159 0.0073 1.0000 1.0000 0.250 -0.0351 0.02165 0.01199 0.0093 1.0000 1.0000 0.500 -0.0186 0.02231 0.01256 0.0089 0.9969 1.0000 0.750 0.0622 0.02335 0.01349 -0.0028 0.9721 1.0000 1.000 0.1358 0.02402 0.01413 -0.0125 0.9452 1.0000 1.250 0.2075 0.02441 0.01452 -0.0213 0.9182 1.0000 1.500 0.2950 0.02432 0.01451 -0.0319 0.8922 1.0000 1.750 0.3604 0.02396 0.01423 -0.0378 0.8629 1.0000 2.000 0.4084 0.02362 0.01394 -0.0400 0.8326 1.0000 2.250 0.4455 0.02338 0.01371 -0.0401 0.8023 1.0000 2.500 0.4756 0.02324 0.01356 -0.0389 0.7722 1.0000 2.750 0.5020 0.02316 0.01344 -0.0370 0.7423 1.0000 3.000 0.5266 0.02314 0.01336 -0.0347 0.7128 1.0000 3.250 0.5502 0.02317 0.01328 -0.0323 0.6834 1.0000 3.500 0.5707 0.02343 0.01346 -0.0298 0.6522 1.0000 3.750 0.5911 0.02377 0.01371 -0.0274 0.6215 1.0000 4.000 0.6117 0.02417 0.01403 -0.0252 0.5919 1.0000 4.250 0.6328 0.02463 0.01439 -0.0232 0.5640 1.0000 4.500 0.6550 0.02507 0.01468 -0.0212 0.5379 1.0000 5.000 0.6966 0.02640 0.01586 -0.0177 0.4870 1.0000 5.250 0.7163 0.02722 0.01668 -0.0160 0.4626 1.0000 5.500 0.7388 0.02791 0.01722 -0.0145 0.4411 1.0000 5.750 0.7578 0.02890 0.01830 -0.0129 0.4192 1.0000 6.000 0.7781 0.02986 0.01927 -0.0115 0.3993 1.0000 6.250 0.7991 0.03086 0.02026 -0.0101 0.3813 1.0000 6.500 0.8197 0.03201 0.02143 -0.0089 0.3648 1.0000 6.750 0.8395 0.03327 0.02278 -0.0076 0.3492 1.0000 7.000 0.8579 0.03482 0.02445 -0.0063 0.3355 1.0000 7.250 0.8761 0.03650 0.02626 -0.0051 0.3235 1.0000 7.500 0.8978 0.03786 0.02761 -0.0041 0.3120 1.0000 7.750 0.9123 0.03982 0.02985 -0.0028 0.3010 1.0000 8.000 0.9240 0.04238 0.03270 -0.0014 0.2928 1.0000 8.250 0.9424 0.04423 0.03465 -0.0004 0.2841 1.0000 8.500 0.9453 0.04757 0.03835 0.0012 0.2769 1.0000 8.750 0.9567 0.05006 0.04104 0.0024 0.2694 1.0000 9.000 0.9623 0.05352 0.04472 0.0036 0.2644 1.0000 9.250 0.9363 0.05978 0.05137 0.0048 0.2622 1.0000 9.500 0.8967 0.06765 0.05944 0.0047 0.2625 1.0000 9.750 0.8528 0.07644 0.06825 0.0030 0.2647 1.0000 10.000 0.8231 0.08543 0.07723 -0.0010 0.2665 1.0000