GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 598 AIRFOIL (goe598-il) Reynolds number: 500,000 Max Cl/Cd: 66.9 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe598-il-500000.txt Download as CSV file: xf-goe598-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 598 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6285 0.07673 0.07466 -0.0110 1.0000 0.0159
-8.000 -0.6276 0.06874 0.06661 -0.0190 1.0000 0.0161
-7.750 -0.6249 0.06221 0.05998 -0.0236 1.0000 0.0164
-7.500 -0.6180 0.05727 0.05493 -0.0262 1.0000 0.0167
-7.250 -0.6082 0.05287 0.05041 -0.0281 1.0000 0.0170
-7.000 -0.5961 0.04857 0.04595 -0.0296 1.0000 0.0174
-6.750 -0.5822 0.04433 0.04152 -0.0307 1.0000 0.0180
-6.500 -0.5664 0.04012 0.03708 -0.0314 1.0000 0.0187
-6.250 -0.5516 0.03024 0.02642 -0.0306 1.0000 0.0145
-6.000 -0.5356 0.02452 0.02010 -0.0298 1.0000 0.0137
-5.750 -0.5141 0.02120 0.01629 -0.0289 1.0000 0.0140
-5.500 -0.4907 0.01903 0.01376 -0.0281 1.0000 0.0145
-5.250 -0.4664 0.01785 0.01225 -0.0274 1.0000 0.0159
-5.000 -0.4440 0.01503 0.00917 -0.0265 1.0000 0.0169
-4.750 -0.4200 0.01385 0.00790 -0.0258 1.0000 0.0182
-4.500 -0.3957 0.01299 0.00696 -0.0251 1.0000 0.0199
-4.250 -0.3713 0.01221 0.00609 -0.0243 1.0000 0.0219
-4.000 -0.3463 0.01216 0.00599 -0.0237 1.0000 0.0242
-3.750 -0.3233 0.01050 0.00419 -0.0229 1.0000 0.0280
-3.500 -0.2990 0.00988 0.00352 -0.0223 1.0000 0.0317
-3.250 -0.2743 0.00949 0.00308 -0.0217 1.0000 0.0353
-3.000 -0.2494 0.00907 0.00259 -0.0211 1.0000 0.0389
-2.750 -0.2244 0.00876 0.00227 -0.0206 1.0000 0.0448
-2.500 -0.1994 0.00850 0.00209 -0.0202 1.0000 0.0599
-2.250 -0.1668 0.00831 0.00194 -0.0214 0.9980 0.0844
-2.000 -0.1285 0.00808 0.00178 -0.0239 0.9946 0.1064
-1.750 -0.0925 0.00752 0.00163 -0.0261 0.9902 0.2139
-1.500 -0.0561 0.00675 0.00148 -0.0286 0.9860 0.3917
-1.250 -0.0186 0.00611 0.00144 -0.0311 0.9822 0.5621
-1.000 0.0148 0.00573 0.00139 -0.0324 0.9727 0.6556
-0.750 0.0461 0.00536 0.00136 -0.0330 0.9608 0.7459
-0.500 0.0718 0.00501 0.00136 -0.0320 0.9454 0.8481
-0.250 0.0961 0.00484 0.00136 -0.0304 0.9278 0.9295
0.000 0.1376 0.00480 0.00131 -0.0330 0.9082 0.9884
0.250 0.1691 0.00481 0.00122 -0.0336 0.8780 1.0000
0.500 0.1934 0.00490 0.00117 -0.0326 0.8486 1.0000
0.750 0.2188 0.00500 0.00116 -0.0318 0.8237 1.0000
1.000 0.2445 0.00511 0.00117 -0.0313 0.8016 1.0000
1.250 0.2705 0.00523 0.00119 -0.0308 0.7782 1.0000
1.500 0.2967 0.00534 0.00124 -0.0303 0.7529 1.0000
1.750 0.3232 0.00545 0.00129 -0.0300 0.7309 1.0000
2.000 0.3496 0.00559 0.00134 -0.0296 0.7041 1.0000
2.250 0.3755 0.00577 0.00140 -0.0291 0.6667 1.0000
2.500 0.4014 0.00600 0.00150 -0.0286 0.6203 1.0000
2.750 0.4261 0.00641 0.00158 -0.0280 0.5231 1.0000
3.000 0.4478 0.00757 0.00188 -0.0274 0.3159 1.0000
3.250 0.4708 0.00869 0.00228 -0.0271 0.1466 1.0000
3.500 0.4943 0.00986 0.00292 -0.0266 0.0280 1.0000
3.750 0.5207 0.01031 0.00342 -0.0263 0.0225 1.0000
4.000 0.5453 0.01132 0.00456 -0.0255 0.0190 1.0000
4.250 0.5708 0.01200 0.00532 -0.0250 0.0182 1.0000
4.500 0.5964 0.01258 0.00595 -0.0246 0.0163 1.0000
4.750 0.6213 0.01346 0.00690 -0.0240 0.0152 1.0000
5.000 0.6459 0.01451 0.00803 -0.0233 0.0142 1.0000
5.250 0.6703 0.01588 0.00951 -0.0225 0.0135 1.0000
5.500 0.6950 0.01807 0.01193 -0.0213 0.0140 1.0000
8.750 0.7707 0.05674 0.05457 -0.0016 0.0131 1.0000
9.000 0.7394 0.06363 0.06160 -0.0040 0.0135 1.0000
9.250 0.7024 0.07616 0.07418 -0.0137 0.0145 1.0000
9.500 0.6865 0.08414 0.08214 -0.0180 0.0148 1.0000
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Polar data table (+)
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