GOE 416A AIRFOIL (goe416a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 416A AIRFOIL (goe416a-il) Reynolds number: 50,000 Max Cl/Cd: 31.06 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe416a-il-50000.txt Download as CSV file: xf-goe416a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 416A AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6126 0.08710 0.08070 -0.0170 1.0000 0.1464
-8.750 -0.6619 0.07654 0.07022 -0.0297 1.0000 0.1429
-8.500 -0.6703 0.06962 0.06319 -0.0336 1.0000 0.1327
-8.250 -0.6888 0.06276 0.05590 -0.0381 1.0000 0.1249
-8.000 -0.6955 0.05652 0.04903 -0.0405 1.0000 0.1159
-7.750 -0.6816 0.05201 0.04420 -0.0406 1.0000 0.1093
-7.500 -0.6732 0.04810 0.03915 -0.0407 1.0000 0.1010
-7.250 -0.6516 0.04401 0.03494 -0.0401 1.0000 0.0978
-7.000 -0.6302 0.04057 0.03114 -0.0394 1.0000 0.0943
-6.750 -0.6072 0.03765 0.02771 -0.0385 1.0000 0.0913
-6.500 -0.5830 0.03520 0.02495 -0.0375 1.0000 0.0914
-6.250 -0.5578 0.03306 0.02261 -0.0364 1.0000 0.0932
-6.000 -0.5319 0.03122 0.02066 -0.0349 1.0000 0.0947
-5.750 -0.5067 0.02970 0.01908 -0.0329 1.0000 0.0958
-5.500 -0.4842 0.02842 0.01774 -0.0308 1.0000 0.0978
-5.250 -0.4638 0.02728 0.01649 -0.0288 1.0000 0.1007
-5.000 -0.4468 0.02575 0.01508 -0.0272 1.0000 0.1066
-4.750 -0.4281 0.02426 0.01358 -0.0263 1.0000 0.1183
-4.500 -0.4082 0.02236 0.01196 -0.0260 1.0000 0.1447
-4.250 -0.3954 0.02086 0.01245 -0.0238 1.0000 0.4678
-4.000 -0.3827 0.02282 0.01445 -0.0172 1.0000 0.5587
-3.750 -0.3758 0.02449 0.01625 -0.0087 1.0000 0.6192
-3.500 -0.3691 0.02536 0.01716 -0.0008 1.0000 0.6615
-3.250 -0.3626 0.02577 0.01753 0.0065 1.0000 0.6978
-3.000 -0.3595 0.02604 0.01782 0.0146 1.0000 0.7382
-2.750 -0.3550 0.02607 0.01784 0.0225 1.0000 0.7725
-2.500 -0.3416 0.02553 0.01717 0.0254 1.0000 0.7932
-2.250 -0.3237 0.02487 0.01635 0.0265 1.0000 0.8052
-2.000 -0.3052 0.02425 0.01557 0.0273 1.0000 0.8167
-1.750 -0.2861 0.02367 0.01486 0.0275 1.0000 0.8277
-1.500 -0.2660 0.02317 0.01423 0.0273 1.0000 0.8382
-1.250 -0.2453 0.02275 0.01372 0.0273 1.0000 0.8475
-1.000 -0.2235 0.02240 0.01326 0.0267 1.0000 0.8569
-0.750 -0.2012 0.02215 0.01292 0.0257 1.0000 0.8665
-0.500 -0.1792 0.02196 0.01269 0.0250 1.0000 0.8759
-0.250 -0.1564 0.02186 0.01255 0.0239 1.0000 0.8861
0.000 -0.1333 0.02186 0.01250 0.0225 1.0000 0.8973
0.250 -0.1097 0.02194 0.01257 0.0210 1.0000 0.9095
0.500 -0.0845 0.02212 0.01278 0.0190 1.0000 0.9237
0.750 -0.0512 0.02250 0.01320 0.0152 0.9974 0.9403
1.000 0.0218 0.02334 0.01404 0.0041 0.9778 0.9568
1.250 0.0984 0.02413 0.01487 -0.0075 0.9579 0.9758
1.500 0.1686 0.02474 0.01555 -0.0178 0.9369 1.0000
1.750 0.2198 0.02519 0.01605 -0.0247 0.9158 1.0000
2.000 0.2782 0.02555 0.01650 -0.0321 0.8948 1.0000
2.250 0.3401 0.02556 0.01661 -0.0389 0.8700 1.0000
2.500 0.3894 0.02542 0.01658 -0.0430 0.8437 1.0000
2.750 0.4355 0.02530 0.01654 -0.0463 0.8212 1.0000
3.000 0.4740 0.02541 0.01670 -0.0485 0.7998 1.0000
3.250 0.5143 0.02537 0.01675 -0.0502 0.7805 1.0000
3.500 0.5445 0.02568 0.01712 -0.0507 0.7579 1.0000
3.750 0.5798 0.02560 0.01710 -0.0508 0.7382 1.0000
4.000 0.6067 0.02598 0.01754 -0.0503 0.7143 1.0000
4.250 0.6364 0.02608 0.01774 -0.0493 0.6917 1.0000
4.500 0.6644 0.02627 0.01799 -0.0480 0.6674 1.0000
4.750 0.6900 0.02668 0.01848 -0.0466 0.6406 1.0000
5.000 0.7156 0.02695 0.01882 -0.0448 0.6118 1.0000
5.250 0.7409 0.02705 0.01893 -0.0426 0.5809 1.0000
5.500 0.7662 0.02714 0.01900 -0.0404 0.5499 1.0000
5.750 0.7906 0.02749 0.01938 -0.0387 0.5197 1.0000
6.000 0.8137 0.02791 0.01991 -0.0370 0.4876 1.0000
6.250 0.8375 0.02789 0.01982 -0.0347 0.4513 1.0000
6.500 0.8587 0.02801 0.01984 -0.0323 0.4090 1.0000
6.750 0.8790 0.02830 0.01993 -0.0298 0.3616 1.0000
7.000 0.8954 0.02883 0.02026 -0.0272 0.3057 1.0000
7.250 0.9118 0.03036 0.02130 -0.0248 0.2490 1.0000
7.500 0.9284 0.03279 0.02351 -0.0230 0.2033 1.0000
7.750 0.9459 0.03509 0.02584 -0.0214 0.1712 1.0000
8.000 0.9645 0.03720 0.02789 -0.0200 0.1476 1.0000
8.250 0.9833 0.03979 0.03064 -0.0187 0.1305 1.0000
8.500 1.0011 0.04287 0.03392 -0.0174 0.1185 1.0000
8.750 1.0193 0.04625 0.03738 -0.0165 0.1094 1.0000
9.000 1.0278 0.05056 0.04238 -0.0147 0.1061 1.0000
9.250 1.0340 0.05471 0.04703 -0.0132 0.1026 1.0000
9.500 1.0492 0.05871 0.05094 -0.0126 0.0972 1.0000
9.750 1.0423 0.06328 0.05607 -0.0109 0.0964 1.0000
10.000 1.0322 0.06814 0.06134 -0.0096 0.0962 1.0000
10.250 1.0198 0.07316 0.06665 -0.0088 0.0966 1.0000
10.500 1.0067 0.07829 0.07198 -0.0083 0.0971 1.0000
10.750 0.9217 0.08824 0.08230 -0.0140 0.1080 1.0000
11.000 0.9039 0.09560 0.08967 -0.0178 0.1098 1.0000
11.250 0.8931 0.10274 0.09680 -0.0209 0.1108 1.0000
11.500 0.8141 0.12665 0.12046 -0.0420 0.1347 1.0000
11.750 0.8010 0.13623 0.12995 -0.0474 0.1485 1.0000
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Polar data table (+)
Polar graphs
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