GOE 416A AIRFOIL (goe416a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 416A AIRFOIL (goe416a-il) Reynolds number: 50,000 Max Cl/Cd: 34.02 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe416a-il-50000-n5.txt Download as CSV file: xf-goe416a-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5982 0.09525 0.08853 -0.0202 1.0000 0.0409
-10.000 -0.6291 0.08626 0.07944 -0.0299 1.0000 0.0396
-9.750 -0.6495 0.08121 0.07420 -0.0354 1.0000 0.0388
-9.500 -0.6525 0.07673 0.06971 -0.0369 1.0000 0.0384
-9.250 -0.6583 0.07254 0.06545 -0.0384 1.0000 0.0380
-9.000 -0.6624 0.06837 0.06114 -0.0396 1.0000 0.0375
-8.750 -0.6643 0.06430 0.05688 -0.0406 1.0000 0.0370
-8.500 -0.6633 0.06035 0.05268 -0.0412 1.0000 0.0365
-8.250 -0.6589 0.05654 0.04857 -0.0415 1.0000 0.0359
-8.000 -0.6509 0.05291 0.04461 -0.0414 1.0000 0.0354
-7.750 -0.6394 0.04949 0.04082 -0.0412 1.0000 0.0348
-7.500 -0.6249 0.04630 0.03717 -0.0407 1.0000 0.0344
-7.250 -0.6078 0.04338 0.03387 -0.0400 1.0000 0.0341
-7.000 -0.5884 0.04073 0.03086 -0.0392 1.0000 0.0339
-6.750 -0.5674 0.03838 0.02822 -0.0382 1.0000 0.0343
-6.500 -0.5455 0.03637 0.02595 -0.0372 1.0000 0.0356
-6.250 -0.5227 0.03462 0.02398 -0.0359 1.0000 0.0369
-6.000 -0.4996 0.03311 0.02230 -0.0343 1.0000 0.0379
-5.750 -0.4777 0.03178 0.02086 -0.0325 1.0000 0.0384
-5.500 -0.4580 0.03057 0.01957 -0.0305 1.0000 0.0391
-5.250 -0.4418 0.02912 0.01804 -0.0288 1.0000 0.0402
-5.000 -0.4253 0.02777 0.01663 -0.0275 1.0000 0.0427
-4.750 -0.4071 0.02658 0.01532 -0.0266 1.0000 0.0464
-4.500 -0.3873 0.02552 0.01404 -0.0258 1.0000 0.0498
-4.250 -0.3672 0.02421 0.01262 -0.0253 1.0000 0.0538
-4.000 -0.3454 0.02313 0.01140 -0.0249 1.0000 0.0617
-3.750 -0.3285 0.01995 0.01030 -0.0259 1.0000 0.3440
-3.500 -0.3094 0.02051 0.01097 -0.0236 1.0000 0.4921
-3.250 -0.2915 0.02117 0.01156 -0.0208 1.0000 0.5470
-3.000 -0.2766 0.02177 0.01223 -0.0171 1.0000 0.5887
-2.750 -0.2617 0.02211 0.01260 -0.0136 1.0000 0.6215
-2.500 -0.2442 0.02215 0.01256 -0.0112 1.0000 0.6414
-2.250 -0.2208 0.02194 0.01212 -0.0113 1.0000 0.6497
-2.000 -0.1987 0.02178 0.01182 -0.0109 1.0000 0.6561
-1.750 -0.1616 0.02166 0.01145 -0.0137 0.9922 0.6635
-1.500 -0.1165 0.02157 0.01118 -0.0177 0.9782 0.6696
-1.250 -0.0713 0.02147 0.01089 -0.0219 0.9653 0.6766
-1.000 -0.0295 0.02137 0.01067 -0.0250 0.9523 0.6821
-0.750 0.0121 0.02127 0.01047 -0.0282 0.9389 0.6884
-0.500 0.0528 0.02116 0.01028 -0.0311 0.9254 0.6942
-0.250 0.0910 0.02105 0.01014 -0.0334 0.9113 0.7001
0.000 0.1285 0.02096 0.01000 -0.0355 0.8967 0.7067
0.250 0.1635 0.02084 0.00990 -0.0369 0.8817 0.7124
0.500 0.1973 0.02076 0.00981 -0.0381 0.8660 0.7193
0.750 0.2282 0.02067 0.00978 -0.0386 0.8492 0.7258
1.000 0.2592 0.02061 0.00975 -0.0392 0.8324 0.7333
1.250 0.2893 0.02053 0.00974 -0.0394 0.8159 0.7407
1.500 0.3192 0.02047 0.00974 -0.0396 0.7994 0.7494
1.750 0.3479 0.02037 0.00974 -0.0393 0.7826 0.7584
2.000 0.3752 0.02030 0.00975 -0.0389 0.7636 0.7688
2.250 0.4019 0.02021 0.00977 -0.0382 0.7431 0.7807
2.500 0.4285 0.02006 0.00969 -0.0371 0.7223 0.7946
2.750 0.4531 0.01994 0.00969 -0.0358 0.6984 0.8118
3.000 0.4788 0.01983 0.00972 -0.0347 0.6768 0.8361
3.250 0.5072 0.01979 0.00987 -0.0344 0.6565 0.8754
3.500 0.5446 0.01978 0.01003 -0.0360 0.6368 1.0000
3.750 0.5737 0.02010 0.01035 -0.0366 0.6173 1.0000
4.000 0.6019 0.02044 0.01072 -0.0369 0.5976 1.0000
4.250 0.6299 0.02078 0.01108 -0.0370 0.5784 1.0000
4.500 0.6568 0.02118 0.01158 -0.0369 0.5580 1.0000
4.750 0.6833 0.02156 0.01205 -0.0367 0.5370 1.0000
5.000 0.7094 0.02196 0.01253 -0.0363 0.5153 1.0000
5.250 0.7348 0.02233 0.01300 -0.0357 0.4913 1.0000
5.500 0.7588 0.02271 0.01340 -0.0348 0.4628 1.0000
5.750 0.7822 0.02316 0.01385 -0.0339 0.4330 1.0000
6.000 0.8055 0.02373 0.01448 -0.0331 0.4046 1.0000
6.250 0.8284 0.02436 0.01521 -0.0323 0.3757 1.0000
6.500 0.8495 0.02497 0.01596 -0.0313 0.3416 1.0000
6.750 0.8691 0.02561 0.01663 -0.0301 0.3039 1.0000
7.000 0.8878 0.02646 0.01743 -0.0290 0.2677 1.0000
7.250 0.9050 0.02762 0.01849 -0.0278 0.2330 1.0000
7.500 0.9210 0.02904 0.01987 -0.0266 0.1976 1.0000
7.750 0.9368 0.03055 0.02146 -0.0253 0.1627 1.0000
8.000 0.9491 0.03221 0.02293 -0.0240 0.1311 1.0000
8.250 0.9586 0.03429 0.02483 -0.0226 0.1050 1.0000
8.500 0.9682 0.03648 0.02704 -0.0211 0.0826 1.0000
8.750 0.9772 0.03876 0.02936 -0.0195 0.0684 1.0000
9.000 0.9864 0.04112 0.03194 -0.0179 0.0584 1.0000
9.250 0.9932 0.04356 0.03434 -0.0164 0.0518 1.0000
9.500 1.0065 0.04622 0.03735 -0.0148 0.0468 1.0000
9.750 1.0142 0.04870 0.04000 -0.0133 0.0427 1.0000
10.000 1.0195 0.05145 0.04273 -0.0119 0.0397 1.0000
10.250 1.0244 0.05462 0.04636 -0.0104 0.0376 1.0000
10.500 1.0256 0.05816 0.05026 -0.0092 0.0364 1.0000
10.750 1.0219 0.06194 0.05437 -0.0084 0.0355 1.0000
11.000 1.0142 0.06602 0.05873 -0.0083 0.0348 1.0000
11.250 1.0032 0.07055 0.06352 -0.0089 0.0343 1.0000
11.500 0.9891 0.07574 0.06897 -0.0107 0.0341 1.0000
11.750 0.9712 0.08191 0.07538 -0.0137 0.0342 1.0000
12.000 0.9498 0.08933 0.08302 -0.0181 0.0347 1.0000
12.250 0.9265 0.09790 0.09177 -0.0236 0.0354 1.0000
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Polar data table (+)
Polar graphs
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