GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: GOE 598 AIRFOIL (goe598-il) Reynolds number: 200,000 Max Cl/Cd: 54.82 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe598-il-200000.txt Download as CSV file: xf-goe598-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6151 0.09069 0.08736 -0.0092 1.0000 0.0360
-8.250 -0.6129 0.08462 0.08124 -0.0161 1.0000 0.0361
-8.000 -0.6198 0.07835 0.07506 -0.0154 1.0000 0.0374
-7.750 -0.6131 0.07557 0.07228 -0.0146 1.0000 0.0387
-7.500 -0.6059 0.07133 0.06801 -0.0173 1.0000 0.0402
-7.250 -0.5977 0.06634 0.06294 -0.0212 1.0000 0.0420
-7.000 -0.5879 0.06065 0.05710 -0.0256 1.0000 0.0446
-5.750 -0.5148 0.04049 0.03587 -0.0320 1.0000 0.0758
-5.500 -0.4950 0.03774 0.03256 -0.0321 1.0000 0.0870
-5.250 -0.4681 0.02523 0.01862 -0.0296 1.0000 0.0374
-5.000 -0.4424 0.02340 0.01645 -0.0286 1.0000 0.0358
-4.750 -0.4184 0.02069 0.01330 -0.0277 1.0000 0.0364
-4.500 -0.3951 0.01779 0.01016 -0.0271 1.0000 0.0400
-4.250 -0.3702 0.01616 0.00836 -0.0261 1.0000 0.0418
-4.000 -0.3455 0.01487 0.00696 -0.0252 1.0000 0.0449
-3.750 -0.3216 0.01377 0.00579 -0.0245 1.0000 0.0508
-3.500 -0.2970 0.01308 0.00509 -0.0237 1.0000 0.0566
-3.250 -0.2725 0.01243 0.00446 -0.0231 1.0000 0.0647
-3.000 -0.2477 0.01189 0.00394 -0.0225 1.0000 0.0790
-2.750 -0.2229 0.01107 0.00325 -0.0221 1.0000 0.1170
-2.500 -0.1988 0.01006 0.00285 -0.0220 1.0000 0.2581
-2.250 -0.1762 0.00887 0.00273 -0.0217 1.0000 0.5230
-2.000 -0.1534 0.00835 0.00268 -0.0206 1.0000 0.6513
-1.750 -0.1324 0.00794 0.00269 -0.0188 1.0000 0.7651
-1.500 -0.1105 0.00770 0.00274 -0.0165 1.0000 0.8980
-1.250 -0.0643 0.00765 0.00265 -0.0203 1.0000 1.0000
-1.000 -0.0427 0.00773 0.00263 -0.0195 1.0000 1.0000
-0.750 -0.0209 0.00784 0.00267 -0.0188 1.0000 1.0000
-0.500 0.0009 0.00801 0.00277 -0.0183 1.0000 1.0000
-0.250 0.0352 0.00816 0.00285 -0.0203 0.9961 1.0000
0.000 0.0804 0.00824 0.00289 -0.0244 0.9877 1.0000
0.250 0.1274 0.00829 0.00292 -0.0287 0.9791 1.0000
0.500 0.1760 0.00827 0.00292 -0.0332 0.9702 1.0000
0.750 0.2191 0.00823 0.00291 -0.0364 0.9583 1.0000
1.000 0.2598 0.00814 0.00285 -0.0388 0.9431 1.0000
1.250 0.2922 0.00805 0.00278 -0.0393 0.9210 1.0000
1.500 0.3206 0.00799 0.00275 -0.0388 0.8997 1.0000
1.750 0.3461 0.00801 0.00278 -0.0378 0.8781 1.0000
2.000 0.3710 0.00805 0.00282 -0.0366 0.8571 1.0000
2.250 0.3946 0.00812 0.00286 -0.0351 0.8313 1.0000
2.500 0.4180 0.00823 0.00297 -0.0336 0.8004 1.0000
2.750 0.4418 0.00836 0.00306 -0.0321 0.7667 1.0000
3.000 0.4640 0.00853 0.00309 -0.0302 0.7104 1.0000
3.250 0.4857 0.00886 0.00313 -0.0283 0.6265 1.0000
3.500 0.5002 0.01036 0.00331 -0.0258 0.3303 1.0000
3.750 0.5155 0.01321 0.00470 -0.0245 0.0516 1.0000
4.000 0.5384 0.01453 0.00606 -0.0235 0.0408 1.0000
4.250 0.5636 0.01528 0.00689 -0.0228 0.0359 1.0000
4.500 0.5876 0.01646 0.00810 -0.0220 0.0331 1.0000
4.750 0.6119 0.01793 0.00960 -0.0211 0.0317 1.0000
5.000 0.6373 0.01971 0.01147 -0.0202 0.0312 1.0000
5.250 0.6614 0.02290 0.01489 -0.0194 0.0288 1.0000
5.500 0.6869 0.02482 0.01713 -0.0185 0.0284 1.0000
5.750 0.7122 0.02743 0.01998 -0.0176 0.0298 1.0000
6.000 0.7474 0.03628 0.03059 -0.0123 0.0837 1.0000
6.250 0.7649 0.03903 0.03361 -0.0116 0.0729 1.0000
6.500 0.7842 0.04196 0.03657 -0.0111 0.0688 1.0000
6.750 0.7964 0.04552 0.04067 -0.0100 0.0601 1.0000
7.000 0.8121 0.04859 0.04384 -0.0095 0.0567 1.0000
9.500 0.6618 0.09353 0.09038 -0.0193 0.0483 1.0000
9.750 0.6542 0.09892 0.09576 -0.0214 0.0473 1.0000
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Polar data table (+)
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