GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 598 AIRFOIL (goe598-il) Reynolds number: 1,000,000 Max Cl/Cd: 72.51 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe598-il-1000000.txt Download as CSV file: xf-goe598-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6526 0.13304 0.13142 0.0270 1.0000 0.0067
-11.000 -0.6484 0.12947 0.12785 0.0256 1.0000 0.0070
-7.000 -0.6411 0.01990 0.01610 -0.0293 1.0000 0.0064
-6.750 -0.6208 0.01673 0.01246 -0.0285 1.0000 0.0066
-6.500 -0.5965 0.01560 0.01116 -0.0281 1.0000 0.0070
-6.250 -0.5713 0.01492 0.01037 -0.0278 1.0000 0.0077
-6.000 -0.5465 0.01404 0.00936 -0.0273 1.0000 0.0083
-5.750 -0.5219 0.01305 0.00822 -0.0268 1.0000 0.0089
-5.500 -0.4973 0.01226 0.00731 -0.0262 1.0000 0.0093
-5.250 -0.4724 0.01179 0.00677 -0.0256 1.0000 0.0097
-5.000 -0.4500 0.01015 0.00494 -0.0247 1.0000 0.0115
-4.750 -0.4248 0.01000 0.00479 -0.0242 1.0000 0.0128
-4.500 -0.4002 0.00967 0.00443 -0.0236 1.0000 0.0142
-4.250 -0.3755 0.00947 0.00420 -0.0229 1.0000 0.0152
-4.000 -0.3517 0.00862 0.00326 -0.0222 1.0000 0.0180
-3.750 -0.3259 0.00843 0.00307 -0.0219 0.9998 0.0208
-3.500 -0.2909 0.00825 0.00289 -0.0236 0.9977 0.0236
-3.250 -0.2561 0.00794 0.00253 -0.0253 0.9956 0.0256
-3.000 -0.2215 0.00747 0.00200 -0.0268 0.9935 0.0304
-2.750 -0.1876 0.00722 0.00169 -0.0283 0.9901 0.0339
-2.500 -0.1529 0.00704 0.00147 -0.0298 0.9864 0.0358
-2.250 -0.1176 0.00688 0.00131 -0.0315 0.9824 0.0386
-2.000 -0.0869 0.00673 0.00119 -0.0321 0.9729 0.0458
-1.750 -0.0571 0.00654 0.00110 -0.0325 0.9608 0.0757
-1.500 -0.0300 0.00642 0.00101 -0.0322 0.9444 0.0904
-1.250 -0.0044 0.00624 0.00091 -0.0316 0.9240 0.1225
-1.000 0.0204 0.00588 0.00081 -0.0310 0.8988 0.2232
-0.750 0.0453 0.00544 0.00072 -0.0305 0.8714 0.3577
-0.500 0.0702 0.00504 0.00067 -0.0301 0.8401 0.5068
-0.250 0.0961 0.00489 0.00066 -0.0297 0.8095 0.5894
0.000 0.1226 0.00475 0.00066 -0.0294 0.7870 0.6576
0.250 0.1489 0.00456 0.00068 -0.0291 0.7667 0.7380
0.500 0.1733 0.00432 0.00072 -0.0282 0.7462 0.8383
0.750 0.1939 0.00416 0.00076 -0.0261 0.7197 0.9419
1.000 0.2367 0.00421 0.00076 -0.0294 0.6934 1.0000
1.250 0.2638 0.00430 0.00078 -0.0292 0.6739 1.0000
1.500 0.2909 0.00440 0.00083 -0.0291 0.6544 1.0000
1.750 0.3177 0.00456 0.00087 -0.0289 0.6223 1.0000
2.000 0.3444 0.00475 0.00092 -0.0287 0.5777 1.0000
2.250 0.3691 0.00539 0.00102 -0.0283 0.4316 1.0000
2.500 0.3940 0.00609 0.00125 -0.0281 0.3063 1.0000
2.750 0.4192 0.00678 0.00147 -0.0280 0.1843 1.0000
3.000 0.4452 0.00730 0.00171 -0.0279 0.1114 1.0000
3.250 0.4704 0.00805 0.00206 -0.0276 0.0223 1.0000
3.500 0.4975 0.00834 0.00238 -0.0274 0.0170 1.0000
3.750 0.5244 0.00868 0.00276 -0.0272 0.0145 1.0000
4.000 0.5499 0.00949 0.00372 -0.0267 0.0116 1.0000
4.250 0.5767 0.00979 0.00404 -0.0265 0.0111 1.0000
4.500 0.6029 0.01028 0.00458 -0.0262 0.0104 1.0000
4.750 0.6288 0.01085 0.00521 -0.0258 0.0096 1.0000
5.000 0.6545 0.01144 0.00586 -0.0255 0.0087 1.0000
5.250 0.6796 0.01221 0.00668 -0.0250 0.0080 1.0000
5.500 0.7035 0.01344 0.00801 -0.0242 0.0077 1.0000
5.750 0.7276 0.01492 0.00963 -0.0234 0.0078 1.0000
6.000 0.7513 0.01656 0.01141 -0.0227 0.0075 1.0000
6.250 0.7736 0.01904 0.01414 -0.0217 0.0073 1.0000
6.500 0.7971 0.02074 0.01605 -0.0210 0.0073 1.0000
6.750 0.8197 0.02277 0.01837 -0.0200 0.0075 1.0000
12.500 0.7972 0.15506 0.15347 -0.0605 0.0068 1.0000
12.750 0.8007 0.15962 0.15803 -0.0623 0.0068 1.0000
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