XFOIL Version 6.96 Calculated polar for: BOEING-VERTOL V23010-1.58 AIRFOIL 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 ------ -------- --------- --------- -------- -------- -------- -10.250 -0.5349 0.09810 0.09562 -0.0117 1.0000 0.0264 -10.000 -0.5404 0.09154 0.08908 -0.0170 1.0000 0.0265 -9.750 -0.5433 0.08637 0.08393 -0.0189 1.0000 0.0266 -9.500 -0.5413 0.08320 0.08078 -0.0192 1.0000 0.0268 -9.250 -0.5468 0.07753 0.07512 -0.0240 1.0000 0.0269 -9.000 -0.5576 0.07271 0.07027 -0.0276 1.0000 0.0269 -8.750 -0.5676 0.06941 0.06695 -0.0280 1.0000 0.0271 -8.500 -0.5757 0.06615 0.06364 -0.0276 1.0000 0.0272 -8.250 -0.5776 0.06297 0.06040 -0.0275 1.0000 0.0273 -8.000 -0.5764 0.06003 0.05741 -0.0271 1.0000 0.0276 -7.750 -0.5739 0.05706 0.05436 -0.0266 1.0000 0.0279 -7.500 -0.5693 0.05419 0.05140 -0.0258 1.0000 0.0284 -7.250 -0.5637 0.05123 0.04832 -0.0248 1.0000 0.0291 -7.000 -0.5634 0.04823 0.04463 -0.0213 1.0000 0.0314 -6.750 -0.5601 0.04320 0.03950 -0.0198 1.0000 0.0317 -6.500 -0.5478 0.04072 0.03702 -0.0186 1.0000 0.0320 -6.250 -0.5349 0.03877 0.03505 -0.0171 1.0000 0.0323 -6.000 -0.5221 0.03699 0.03321 -0.0154 1.0000 0.0327 -5.750 -0.5095 0.03532 0.03146 -0.0135 1.0000 0.0333 -5.500 -0.4974 0.03368 0.02972 -0.0113 1.0000 0.0341 -5.250 -0.4704 0.03383 0.02928 -0.0097 0.9979 0.0371 -5.000 -0.4397 0.02863 0.02396 -0.0125 0.9931 0.0380 -4.750 -0.4035 0.02663 0.02195 -0.0152 0.9869 0.0389 -4.500 -0.3666 0.02506 0.02029 -0.0176 0.9803 0.0404 -4.250 -0.3259 0.02617 0.02096 -0.0187 0.9725 0.0441 -4.000 -0.2972 0.02171 0.01641 -0.0202 0.9622 0.0456 -3.750 -0.2633 0.02027 0.01495 -0.0217 0.9501 0.0470 -3.500 -0.2301 0.01930 0.01386 -0.0226 0.9335 0.0496 -3.250 -0.1998 0.01833 0.01255 -0.0225 0.9109 0.0540 -3.000 -0.1721 0.01701 0.01122 -0.0224 0.8848 0.0556 -2.750 -0.1413 0.01491 0.00852 -0.0205 0.8622 0.0375 -2.500 -0.1152 0.01369 0.00726 -0.0198 0.8383 0.0364 -2.250 -0.0901 0.01269 0.00619 -0.0189 0.8119 0.0355 -2.000 -0.0654 0.01200 0.00538 -0.0178 0.7855 0.0350 -1.500 -0.0164 0.01110 0.00426 -0.0158 0.7432 0.0354 -1.250 0.0084 0.01077 0.00385 -0.0149 0.7231 0.0358 -1.000 0.0334 0.01049 0.00349 -0.0141 0.7014 0.0364 -0.750 0.0585 0.01028 0.00319 -0.0133 0.6754 0.0369 -0.500 0.0825 0.00997 0.00278 -0.0123 0.6412 0.0384 -0.250 0.1070 0.00989 0.00256 -0.0115 0.6004 0.0402 0.000 0.1318 0.00987 0.00237 -0.0107 0.5587 0.0418 0.250 0.1566 0.00992 0.00222 -0.0099 0.5105 0.0439 0.500 0.1807 0.00997 0.00205 -0.0091 0.4511 0.0482 0.750 0.2051 0.01009 0.00194 -0.0083 0.3954 0.0616 1.000 0.2098 0.00794 0.00183 -0.0044 0.3672 0.7259 1.250 0.2281 0.00772 0.00198 -0.0019 0.3397 0.8582 1.500 0.2601 0.00785 0.00213 -0.0022 0.3170 0.9269 1.750 0.3054 0.00811 0.00229 -0.0056 0.2998 0.9602 2.000 0.3543 0.00839 0.00246 -0.0097 0.2880 0.9770 2.250 0.4010 0.00860 0.00259 -0.0136 0.2785 0.9863 2.500 0.4429 0.00884 0.00270 -0.0166 0.2701 0.9923 2.750 0.4814 0.00892 0.00276 -0.0188 0.2645 0.9961 3.000 0.5208 0.00905 0.00284 -0.0213 0.2589 0.9997 3.250 0.5435 0.00923 0.00295 -0.0204 0.2542 1.0000 3.500 0.5659 0.00932 0.00304 -0.0193 0.2510 1.0000 3.750 0.5885 0.00942 0.00314 -0.0182 0.2472 1.0000 4.000 0.6112 0.00956 0.00326 -0.0172 0.2432 1.0000 4.250 0.6336 0.00978 0.00342 -0.0161 0.2389 1.0000 4.500 0.6569 0.00992 0.00358 -0.0152 0.2353 1.0000 4.750 0.6807 0.01003 0.00371 -0.0143 0.2314 1.0000 5.000 0.7044 0.01018 0.00384 -0.0134 0.2268 1.0000 5.250 0.7272 0.01046 0.00406 -0.0124 0.2215 1.0000 5.500 0.7519 0.01053 0.00419 -0.0117 0.2176 1.0000 5.750 0.7762 0.01065 0.00433 -0.0109 0.2126 1.0000 6.000 0.7996 0.01090 0.00452 -0.0101 0.2069 1.0000 6.250 0.8244 0.01099 0.00468 -0.0094 0.2011 1.0000 6.500 0.8487 0.01114 0.00480 -0.0087 0.1936 1.0000 6.750 0.8733 0.01128 0.00496 -0.0080 0.1855 1.0000 7.000 0.8971 0.01150 0.00513 -0.0072 0.1767 1.0000 7.250 0.9218 0.01165 0.00531 -0.0066 0.1683 1.0000 7.500 0.9455 0.01189 0.00552 -0.0058 0.1596 1.0000 7.750 0.9692 0.01215 0.00575 -0.0051 0.1512 1.0000 8.000 0.9928 0.01243 0.00604 -0.0044 0.1446 1.0000 8.250 1.0160 0.01275 0.00634 -0.0036 0.1384 1.0000 8.500 1.0393 0.01307 0.00668 -0.0028 0.1329 1.0000 8.750 1.0626 0.01340 0.00702 -0.0021 0.1273 1.0000 9.000 1.0850 0.01380 0.00742 -0.0013 0.1217 1.0000 9.250 1.1086 0.01409 0.00776 -0.0006 0.1164 1.0000 9.500 1.1302 0.01457 0.00822 0.0003 0.1097 1.0000 9.750 1.1541 0.01483 0.00855 0.0009 0.1033 1.0000 10.000 1.1764 0.01524 0.00897 0.0017 0.0944 1.0000 10.250 1.1977 0.01577 0.00942 0.0025 0.0767 1.0000 10.500 1.2141 0.01678 0.01025 0.0039 0.0545 1.0000 10.750 1.2308 0.01772 0.01115 0.0053 0.0460 1.0000 11.000 1.2463 0.01871 0.01214 0.0069 0.0414 1.0000 11.250 1.2645 0.01939 0.01291 0.0081 0.0389 1.0000 11.500 1.2793 0.02030 0.01386 0.0097 0.0365 1.0000 11.750 1.2918 0.02132 0.01495 0.0116 0.0346 1.0000 12.000 1.3065 0.02210 0.01582 0.0132 0.0332 1.0000 12.250 1.3186 0.02299 0.01679 0.0150 0.0319 1.0000 12.500 1.3242 0.02405 0.01791 0.0178 0.0309 1.0000 12.750 1.3207 0.02549 0.01942 0.0214 0.0299 1.0000 13.000 1.3234 0.02674 0.02077 0.0239 0.0292 1.0000 13.250 1.3287 0.02796 0.02210 0.0257 0.0284 1.0000 13.500 1.3318 0.02946 0.02370 0.0272 0.0277 1.0000 13.750 1.3321 0.03134 0.02568 0.0283 0.0269 1.0000 14.000 1.3303 0.03365 0.02809 0.0288 0.0264 1.0000 14.250 1.3259 0.03652 0.03106 0.0287 0.0259 1.0000 14.500 1.3182 0.04008 0.03472 0.0279 0.0255 1.0000 14.750 1.3051 0.04461 0.03936 0.0263 0.0251 1.0000 15.000 1.2930 0.04930 0.04416 0.0244 0.0249 1.0000 15.250 1.2858 0.05366 0.04865 0.0224 0.0246 1.0000 15.500 1.2782 0.05824 0.05337 0.0201 0.0243 1.0000 15.750 1.2693 0.06307 0.05831 0.0178 0.0241 1.0000 16.000 1.2589 0.06816 0.06352 0.0153 0.0238 1.0000 16.250 1.2483 0.07324 0.06871 0.0129 0.0236 1.0000 16.500 1.2381 0.07834 0.07392 0.0105 0.0233 1.0000 16.750 1.2288 0.08334 0.07901 0.0081 0.0231 1.0000