GOE 54 AIRFOIL (goe54-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 54 AIRFOIL (goe54-il) Reynolds number: 500,000 Max Cl/Cd: 87.39 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe54-il-500000.txt Download as CSV file: xf-goe54-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 54 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4835 0.12985 0.12774 0.0371 1.0000 0.0171
-10.500 -0.4846 0.12655 0.12445 0.0354 1.0000 0.0171
-10.250 -0.6068 0.13275 0.13052 0.0539 1.0000 0.0170
-10.000 -0.6024 0.12926 0.12705 0.0514 1.0000 0.0171
-9.750 -0.5991 0.12580 0.12359 0.0486 1.0000 0.0171
-9.250 -0.5883 0.11784 0.11567 0.0467 1.0000 0.0173
-9.000 -0.5818 0.11435 0.11218 0.0464 1.0000 0.0174
-8.750 -0.5759 0.11104 0.10889 0.0454 1.0000 0.0176
-8.500 -0.5700 0.10776 0.10563 0.0441 1.0000 0.0178
-8.250 -0.5641 0.10447 0.10236 0.0426 1.0000 0.0181
-8.000 -0.5582 0.10115 0.09905 0.0408 1.0000 0.0184
-7.750 -0.5524 0.09778 0.09570 0.0389 1.0000 0.0188
-7.500 -0.5461 0.09432 0.09226 0.0364 1.0000 0.0193
-7.250 -0.5341 0.09032 0.08826 0.0316 1.0000 0.0199
-7.000 -0.4390 0.07265 0.07071 0.0111 1.0000 0.0203
-6.750 -0.4316 0.06870 0.06678 0.0131 1.0000 0.0204
-6.500 -0.4761 0.07521 0.07306 0.0093 1.0000 0.0204
-6.250 -0.4638 0.07171 0.06957 0.0090 1.0000 0.0205
-6.000 -0.4472 0.06819 0.06604 0.0070 1.0000 0.0207
-5.750 -0.4271 0.06460 0.06243 0.0040 1.0000 0.0210
-5.500 -0.4042 0.06095 0.05874 0.0007 1.0000 0.0215
-5.250 -0.3792 0.05724 0.05497 -0.0028 1.0000 0.0222
-5.000 -0.3369 0.05303 0.05051 -0.0089 1.0000 0.0236
-4.750 -0.3005 0.04899 0.04616 -0.0126 1.0000 0.0237
-4.500 -0.2858 0.04592 0.04237 -0.0124 0.6991 0.0239
-4.250 -0.2658 0.04344 0.03952 -0.0131 0.5981 0.0241
-4.000 -0.2420 0.04108 0.03686 -0.0142 0.5392 0.0244
-3.750 -0.2156 0.03876 0.03430 -0.0153 0.5115 0.0249
-3.500 -0.1875 0.03649 0.03181 -0.0163 0.4936 0.0257
-3.250 -0.1490 0.03511 0.02996 -0.0165 0.4805 0.0276
-2.750 -0.0934 0.02975 0.02418 -0.0179 0.4590 0.0281
-2.500 -0.0673 0.02779 0.02213 -0.0185 0.4494 0.0284
-2.250 -0.0398 0.02619 0.02038 -0.0190 0.4399 0.0289
-2.000 -0.0112 0.02476 0.01879 -0.0192 0.4312 0.0297
-1.750 0.0185 0.02360 0.01742 -0.0192 0.4222 0.0313
-1.500 0.0525 0.02420 0.01747 -0.0180 0.4134 0.0324
-1.250 0.0794 0.02112 0.01438 -0.0189 0.4042 0.0329
-1.000 0.1074 0.01975 0.01296 -0.0192 0.3936 0.0335
-0.750 0.1359 0.01877 0.01185 -0.0193 0.3829 0.0345
-0.500 0.1652 0.01805 0.01095 -0.0192 0.3711 0.0363
-0.250 0.1959 0.01812 0.01062 -0.0185 0.3570 0.0381
0.000 -0.3200 0.05938 0.05527 0.0620 0.4236 0.0281
0.250 0.2523 0.01577 0.00817 -0.0189 0.3226 0.0404
0.500 0.2812 0.01548 0.00770 -0.0187 0.3061 0.0428
0.750 0.3104 0.01500 0.00699 -0.0185 0.2923 0.0451
1.000 0.3386 0.01427 0.00624 -0.0186 0.2807 0.0469
1.250 0.3675 0.01505 0.00677 -0.0181 0.2717 0.0514
2.750 -0.47831384.765501384.76099 0.1717 0.3077 0.0342
3.000 0.5663 0.01195 0.00372 -0.0174 0.2384 0.0657
3.250 0.5948 0.01217 0.00380 -0.0170 0.2346 0.0519
3.500 0.6232 0.01207 0.00370 -0.0170 0.2305 0.0496
3.750 0.6516 0.01202 0.00370 -0.0169 0.2275 0.0507
4.000 0.6801 0.01197 0.00369 -0.0168 0.2239 0.0506
4.250 0.7084 0.01200 0.00371 -0.0167 0.2200 0.0506
4.500 0.7363 0.01221 0.00387 -0.0166 0.2153 0.0513
4.750 0.7647 0.01217 0.00388 -0.0166 0.2123 0.0536
5.000 0.7929 0.01222 0.00395 -0.0165 0.2081 0.0587
5.250 0.8208 0.01235 0.00407 -0.0163 0.2035 0.0672
5.500 0.8440 0.01087 0.00439 -0.0157 0.1989 1.0000
5.750 0.8721 0.01097 0.00452 -0.0155 0.1943 1.0000
6.000 0.8999 0.01111 0.00464 -0.0154 0.1884 1.0000
6.250 0.9277 0.01130 0.00481 -0.0154 0.1820 1.0000
6.500 0.9555 0.01143 0.00496 -0.0152 0.1760 1.0000
6.750 0.9829 0.01170 0.00519 -0.0152 0.1701 1.0000
7.000 1.0107 0.01183 0.00537 -0.0151 0.1645 1.0000
7.250 1.0381 0.01207 0.00560 -0.0150 0.1591 1.0000
7.500 1.0654 0.01234 0.00590 -0.0149 0.1547 1.0000
7.750 1.0928 0.01258 0.00617 -0.0147 0.1503 1.0000
8.000 1.1198 0.01287 0.00646 -0.0146 0.1454 1.0000
8.250 1.1468 0.01315 0.00680 -0.0145 0.1406 1.0000
8.500 1.1737 0.01343 0.00711 -0.0144 0.1360 1.0000
8.750 1.2000 0.01386 0.00752 -0.0143 0.1306 1.0000
9.000 1.2270 0.01410 0.00787 -0.0142 0.1258 1.0000
9.250 1.2531 0.01452 0.00828 -0.0142 0.1190 1.0000
9.500 1.2795 0.01483 0.00867 -0.0140 0.1123 1.0000
9.750 1.3053 0.01531 0.00915 -0.0140 0.1029 1.0000
10.000 1.3310 0.01579 0.00964 -0.0139 0.0869 1.0000
10.250 1.3541 0.01691 0.01057 -0.0138 0.0626 1.0000
10.500 1.3771 0.01798 0.01165 -0.0137 0.0546 1.0000
10.750 1.4002 0.01897 0.01273 -0.0135 0.0498 1.0000
11.000 1.4230 0.01992 0.01376 -0.0133 0.0463 1.0000
11.250 1.4436 0.02128 0.01519 -0.0131 0.0426 1.0000
11.500 1.4666 0.02203 0.01608 -0.0128 0.0402 1.0000
11.750 1.4874 0.02315 0.01728 -0.0126 0.0375 1.0000
12.000 1.5050 0.02475 0.01899 -0.0124 0.0349 1.0000
12.250 1.5257 0.02570 0.02007 -0.0121 0.0329 1.0000
12.500 1.5438 0.02701 0.02147 -0.0119 0.0308 1.0000
12.750 1.5553 0.02920 0.02378 -0.0117 0.0287 1.0000
13.000 1.5723 0.03047 0.02519 -0.0115 0.0272 1.0000
13.250 1.5853 0.03221 0.02706 -0.0114 0.0258 1.0000
13.500 1.5910 0.03484 0.02981 -0.0118 0.0247 1.0000
13.750 1.5796 0.04014 0.03533 -0.0152 0.0241 1.0000
14.000 1.5580 0.05059 0.04604 -0.0258 0.0241 1.0000
14.250 1.5316 0.05876 0.05436 -0.0308 0.0243 1.0000
14.500 1.5047 0.06646 0.06219 -0.0349 0.0244 1.0000
14.750 1.4789 0.07388 0.06972 -0.0384 0.0244 1.0000
15.000 1.4553 0.08095 0.07689 -0.0417 0.0244 1.0000
15.250 1.4339 0.08770 0.08371 -0.0447 0.0242 1.0000
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Polar data table (+)
Polar graphs
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