XFOIL Version 6.96 Calculated polar for: Dayton-Wright T-1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.000 -0.7269 0.09561 0.09332 -0.0659 1.0000 0.0220 -17.750 -0.8185 0.07694 0.07441 -0.0787 1.0000 0.0220 -17.500 -0.8888 0.06325 0.06056 -0.0886 1.0000 0.0220 -17.250 -1.0259 0.03814 0.03508 -0.1118 1.0000 0.0217 -17.000 -1.0222 0.03298 0.02968 -0.1178 0.9953 0.0218 -16.750 -1.0130 0.03056 0.02711 -0.1192 0.9901 0.0220 -16.500 -0.9990 0.02882 0.02526 -0.1197 0.9845 0.0221 -16.250 -0.9821 0.02743 0.02376 -0.1198 0.9794 0.0223 -16.000 -0.9645 0.02621 0.02243 -0.1195 0.9737 0.0225 -15.750 -0.9461 0.02510 0.02121 -0.1191 0.9688 0.0227 -15.500 -0.9293 0.02407 0.02007 -0.1182 0.9623 0.0229 -15.250 -0.9106 0.02312 0.01901 -0.1173 0.9569 0.0230 -15.000 -0.8925 0.02228 0.01806 -0.1162 0.9503 0.0232 -14.750 -0.8729 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15.000 1.5667 0.06222 0.05651 -0.0389 0.0091 1.0000 15.250 1.5621 0.06586 0.06023 -0.0392 0.0090 1.0000 15.500 1.5560 0.06979 0.06425 -0.0396 0.0088 1.0000 15.750 1.5480 0.07402 0.06857 -0.0402 0.0087 1.0000 16.000 1.5386 0.07850 0.07315 -0.0409 0.0085 1.0000 16.250 1.5290 0.08310 0.07785 -0.0417 0.0084 1.0000 16.500 1.5201 0.08768 0.08254 -0.0427 0.0083 1.0000 16.750 1.5099 0.09251 0.08747 -0.0438 0.0083 1.0000 17.000 1.4988 0.09750 0.09256 -0.0450 0.0082 1.0000 17.250 1.4875 0.10260 0.09776 -0.0464 0.0081 1.0000 17.500 1.4761 0.10778 0.10305 -0.0479 0.0081 1.0000 17.750 1.4647 0.11300 0.10837 -0.0495 0.0080 1.0000