GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: GOE 55 AIRFOIL (goe55-il) Reynolds number: 500,000 Max Cl/Cd: 64.71 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe55-il-500000.txt Download as CSV file: xf-goe55-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 55 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.6342 0.08722 0.08496 0.0184 1.0000 0.0306
-7.500 -0.6289 0.08042 0.07815 0.0136 1.0000 0.0310
-7.250 -0.6173 0.07761 0.07533 0.0133 1.0000 0.0314
-7.000 -0.6029 0.07466 0.07238 0.0118 1.0000 0.0318
-6.750 -0.5864 0.07129 0.06899 0.0091 1.0000 0.0324
-6.500 -0.5676 0.06730 0.06496 0.0052 1.0000 0.0333
-6.250 -0.5442 0.06203 0.05960 -0.0009 1.0000 0.0352
-6.000 -0.5116 0.04995 0.04714 -0.0134 1.0000 0.0377
-5.750 -0.4909 0.04791 0.04508 -0.0140 1.0000 0.0382
-5.500 -0.4679 0.04561 0.04273 -0.0151 1.0000 0.0390
-5.250 -0.4425 0.04276 0.03978 -0.0168 1.0000 0.0402
-5.000 -0.4055 0.03373 0.03000 -0.0223 1.0000 0.0453
-4.750 -0.3811 0.03209 0.02840 -0.0228 1.0000 0.0460
-4.500 -0.3552 0.03067 0.02696 -0.0233 1.0000 0.0470
-4.250 -0.3276 0.02887 0.02505 -0.0239 1.0000 0.0486
-4.000 -0.2941 0.02842 0.02418 -0.0239 1.0000 0.0535
-3.750 -0.2610 0.01843 0.01334 -0.0266 1.0000 0.0424
-3.500 -0.2305 0.01572 0.01025 -0.0270 1.0000 0.0411
-3.250 -0.2008 0.01421 0.00851 -0.0272 1.0000 0.0410
-3.000 -0.1716 0.01317 0.00734 -0.0272 1.0000 0.0413
-2.750 -0.1425 0.01236 0.00645 -0.0272 1.0000 0.0419
-2.500 -0.1135 0.01180 0.00584 -0.0271 1.0000 0.0430
-2.250 -0.0863 0.01163 0.00540 -0.0261 0.8439 0.0437
-2.000 -0.0643 0.01235 0.00495 -0.0245 0.5428 0.0442
-1.500 -0.0068 0.01166 0.00367 -0.0248 0.4162 0.0464
-1.250 0.0221 0.01144 0.00338 -0.0248 0.3995 0.0479
-1.000 0.0510 0.01128 0.00318 -0.0248 0.3881 0.0501
-0.750 0.0799 0.01117 0.00302 -0.0248 0.3797 0.0526
-0.500 0.1090 0.01095 0.00276 -0.0248 0.3718 0.0557
-0.250 0.1380 0.01083 0.00262 -0.0248 0.3639 0.0608
0.000 0.1670 0.01070 0.00251 -0.0248 0.3564 0.0711
0.250 0.1960 0.01064 0.00251 -0.0248 0.3502 0.0989
0.500 0.2249 0.01051 0.00245 -0.0248 0.3438 0.1244
0.750 0.2538 0.01035 0.00247 -0.0250 0.3367 0.1915
1.000 0.2815 0.00877 0.00258 -0.0256 0.3301 0.7450
1.250 0.3033 0.00812 0.00241 -0.0235 0.3242 1.0000
1.500 0.3320 0.00826 0.00246 -0.0234 0.3184 1.0000
1.750 0.3608 0.00832 0.00248 -0.0233 0.3116 1.0000
2.000 0.3894 0.00846 0.00252 -0.0232 0.3036 1.0000
2.250 0.4181 0.00849 0.00256 -0.0231 0.2966 1.0000
2.500 0.4467 0.00861 0.00260 -0.0230 0.2888 1.0000
2.750 0.4754 0.00866 0.00265 -0.0229 0.2806 1.0000
3.000 0.5040 0.00876 0.00271 -0.0228 0.2711 1.0000
3.250 0.5325 0.00884 0.00275 -0.0227 0.2581 1.0000
3.500 0.5611 0.00894 0.00280 -0.0227 0.2418 1.0000
3.750 0.5895 0.00911 0.00287 -0.0226 0.2101 1.0000
4.500 0.6735 0.01050 0.00365 -0.0227 0.0948 1.0000
4.750 0.7013 0.01121 0.00407 -0.0228 0.0523 1.0000
5.000 0.7294 0.01164 0.00447 -0.0228 0.0467 1.0000
5.250 0.7575 0.01201 0.00487 -0.0227 0.0446 1.0000
5.500 0.7855 0.01239 0.00530 -0.0226 0.0430 1.0000
5.750 0.8133 0.01282 0.00576 -0.0225 0.0414 1.0000
6.000 0.8410 0.01332 0.00629 -0.0224 0.0402 1.0000
6.250 0.8684 0.01392 0.00694 -0.0223 0.0391 1.0000
6.500 0.8953 0.01469 0.00776 -0.0222 0.0381 1.0000
6.750 0.9217 0.01557 0.00871 -0.0220 0.0374 1.0000
7.000 0.9486 0.01614 0.00934 -0.0219 0.0370 1.0000
7.250 0.9753 0.01673 0.01000 -0.0217 0.0363 1.0000
7.500 1.0015 0.01743 0.01077 -0.0214 0.0357 1.0000
7.750 1.0274 0.01822 0.01162 -0.0212 0.0351 1.0000
8.000 1.0530 0.01908 0.01256 -0.0209 0.0345 1.0000
8.250 1.0782 0.01998 0.01353 -0.0206 0.0338 1.0000
8.500 1.1031 0.02089 0.01451 -0.0203 0.0330 1.0000
8.750 1.1274 0.02194 0.01560 -0.0200 0.0323 1.0000
9.000 1.1505 0.02336 0.01708 -0.0196 0.0316 1.0000
9.250 1.1715 0.02572 0.01955 -0.0191 0.0309 1.0000
9.500 1.1928 0.02763 0.02167 -0.0186 0.0303 1.0000
9.750 1.2164 0.02833 0.02255 -0.0182 0.0298 1.0000
10.000 1.2384 0.02954 0.02395 -0.0177 0.0291 1.0000
10.250 1.2590 0.03106 0.02566 -0.0172 0.0285 1.0000
10.500 1.2783 0.03271 0.02749 -0.0167 0.0279 1.0000
10.750 1.2973 0.03419 0.02910 -0.0161 0.0272 1.0000
11.000 1.3163 0.03552 0.03051 -0.0156 0.0265 1.0000
11.250 1.3328 0.03723 0.03230 -0.0151 0.0258 1.0000
11.500 1.3423 0.04069 0.03589 -0.0145 0.0252 1.0000
11.750 1.3386 0.04567 0.04128 -0.0136 0.0247 1.0000
12.000 1.3437 0.04760 0.04349 -0.0129 0.0244 1.0000
12.250 1.3404 0.05055 0.04672 -0.0122 0.0242 1.0000
12.500 1.3252 0.05408 0.05047 -0.0114 0.0240 1.0000
12.750 1.3056 0.05974 0.05636 -0.0154 0.0240 1.0000
13.000 1.2839 0.06846 0.06531 -0.0241 0.0240 1.0000
13.250 1.2570 0.07884 0.07589 -0.0332 0.0242 1.0000
13.500 1.2238 0.09007 0.08728 -0.0416 0.0244 1.0000
13.750 1.1842 0.10246 0.09980 -0.0496 0.0246 1.0000
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Polar data table (+)
Polar graphs
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