GOE 416A AIRFOIL (goe416a-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 416A AIRFOIL (goe416a-il) Reynolds number: 200,000 Max Cl/Cd: 57.71 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe416a-il-200000-n5.txt Download as CSV file: xf-goe416a-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5042 0.07784 0.07452 -0.0292 1.0000 0.0193
-10.250 -0.5193 0.07241 0.06900 -0.0329 1.0000 0.0190
-8.750 -0.6651 0.05556 0.05149 -0.0407 1.0000 0.0156
-8.500 -0.6677 0.05086 0.04652 -0.0415 1.0000 0.0142
-8.250 -0.6636 0.04673 0.04208 -0.0416 1.0000 0.0136
-8.000 -0.6538 0.04366 0.03874 -0.0415 1.0000 0.0140
-7.750 -0.6414 0.04074 0.03555 -0.0411 1.0000 0.0145
-7.500 -0.6274 0.03769 0.03214 -0.0404 1.0000 0.0144
-7.250 -0.6117 0.03419 0.02821 -0.0393 1.0000 0.0131
-7.000 -0.5932 0.03125 0.02472 -0.0377 1.0000 0.0119
-6.750 -0.5732 0.02934 0.02252 -0.0365 1.0000 0.0116
-6.500 -0.5533 0.02756 0.02051 -0.0353 1.0000 0.0114
-6.250 -0.5331 0.02580 0.01856 -0.0342 1.0000 0.0114
-6.000 -0.5129 0.02408 0.01669 -0.0329 1.0000 0.0115
-5.750 -0.4928 0.02248 0.01498 -0.0317 1.0000 0.0117
-5.500 -0.4728 0.02106 0.01350 -0.0305 1.0000 0.0119
-5.250 -0.4399 0.01971 0.01212 -0.0321 0.9958 0.0132
-5.000 -0.4052 0.01855 0.01090 -0.0338 0.9904 0.0145
-4.750 -0.3697 0.01731 0.00961 -0.0356 0.9850 0.0150
-4.500 -0.3353 0.01621 0.00846 -0.0373 0.9783 0.0157
-4.250 -0.3019 0.01526 0.00742 -0.0389 0.9702 0.0165
-4.000 -0.2674 0.01462 0.00667 -0.0405 0.9620 0.0182
-3.750 -0.2332 0.01375 0.00559 -0.0421 0.9527 0.0203
-3.500 -0.2007 0.01321 0.00488 -0.0431 0.9412 0.0228
-3.250 -0.1689 0.01282 0.00435 -0.0439 0.9298 0.0284
-2.750 -0.1135 0.01050 0.00363 -0.0454 0.9064 0.4599
-2.500 -0.0844 0.01059 0.00366 -0.0454 0.8934 0.4860
-2.250 -0.0566 0.01084 0.00386 -0.0450 0.8794 0.5181
-2.000 -0.0292 0.01095 0.00394 -0.0445 0.8649 0.5338
-1.750 -0.0016 0.01091 0.00385 -0.0442 0.8495 0.5386
-1.500 0.0260 0.01088 0.00370 -0.0440 0.8318 0.5430
-1.250 0.0534 0.01086 0.00353 -0.0437 0.8115 0.5463
-1.000 0.0806 0.01082 0.00340 -0.0433 0.7888 0.5488
-0.750 0.1078 0.01080 0.00328 -0.0430 0.7663 0.5513
-0.500 0.1352 0.01081 0.00318 -0.0427 0.7437 0.5539
-0.250 0.1626 0.01084 0.00309 -0.0425 0.7227 0.5569
0.000 0.1904 0.01088 0.00302 -0.0424 0.7026 0.5600
0.250 0.2180 0.01093 0.00297 -0.0423 0.6830 0.5625
0.500 0.2454 0.01098 0.00295 -0.0421 0.6634 0.5648
0.750 0.2730 0.01105 0.00296 -0.0420 0.6451 0.5677
1.000 0.3008 0.01112 0.00297 -0.0420 0.6282 0.5710
1.250 0.3286 0.01121 0.00300 -0.0420 0.6107 0.5742
1.500 0.3563 0.01130 0.00304 -0.0420 0.5939 0.5770
1.750 0.3839 0.01137 0.00311 -0.0419 0.5770 0.5797
2.000 0.4116 0.01144 0.00319 -0.0419 0.5584 0.5831
2.250 0.4394 0.01153 0.00328 -0.0419 0.5409 0.5869
2.500 0.4672 0.01163 0.00337 -0.0419 0.5240 0.5908
2.750 0.4946 0.01172 0.00349 -0.0419 0.5063 0.5941
3.000 0.5219 0.01184 0.00362 -0.0418 0.4887 0.5978
3.250 0.5491 0.01199 0.00375 -0.0418 0.4692 0.6021
3.500 0.5761 0.01217 0.00393 -0.0416 0.4495 0.6065
3.750 0.6028 0.01235 0.00413 -0.0415 0.4318 0.6113
4.000 0.6299 0.01253 0.00434 -0.0414 0.4165 0.6170
4.250 0.6568 0.01271 0.00460 -0.0413 0.3988 0.6226
4.500 0.6834 0.01291 0.00485 -0.0412 0.3774 0.6289
4.750 0.7095 0.01316 0.00509 -0.0410 0.3444 0.6358
5.250 0.7593 0.01390 0.00570 -0.0404 0.2628 0.6532
5.500 0.7841 0.01431 0.00611 -0.0401 0.2399 0.6651
5.750 0.8091 0.01465 0.00655 -0.0398 0.2233 0.6809
6.000 0.8338 0.01496 0.00700 -0.0393 0.2080 0.7034
6.250 0.8579 0.01523 0.00751 -0.0388 0.1896 0.7418
6.500 0.8801 0.01525 0.00789 -0.0375 0.1524 0.9388
6.750 0.9015 0.01626 0.00854 -0.0371 0.1039 1.0000
7.000 0.9241 0.01711 0.00926 -0.0367 0.0743 1.0000
7.250 0.9464 0.01796 0.01001 -0.0362 0.0537 1.0000
7.500 0.9685 0.01882 0.01086 -0.0356 0.0385 1.0000
7.750 0.9895 0.01981 0.01181 -0.0349 0.0241 1.0000
8.000 1.0102 0.02082 0.01286 -0.0341 0.0182 1.0000
8.250 1.0296 0.02193 0.01412 -0.0331 0.0153 1.0000
8.500 1.0455 0.02345 0.01578 -0.0316 0.0132 1.0000
8.750 1.0641 0.02450 0.01700 -0.0306 0.0120 1.0000
9.000 1.0811 0.02569 0.01834 -0.0294 0.0109 1.0000
9.250 1.0960 0.02706 0.01986 -0.0280 0.0103 1.0000
9.500 1.1093 0.02854 0.02147 -0.0265 0.0097 1.0000
9.750 1.1203 0.03018 0.02326 -0.0248 0.0093 1.0000
10.000 1.1281 0.03210 0.02533 -0.0229 0.0089 1.0000
10.250 1.1308 0.03432 0.02772 -0.0203 0.0086 1.0000
10.500 1.1321 0.03683 0.03040 -0.0179 0.0083 1.0000
10.750 1.1377 0.03869 0.03248 -0.0163 0.0080 1.0000
11.000 1.1412 0.04080 0.03480 -0.0148 0.0077 1.0000
11.250 1.1428 0.04316 0.03737 -0.0137 0.0074 1.0000
11.500 1.1420 0.04589 0.04031 -0.0129 0.0071 1.0000
11.750 1.1392 0.04901 0.04368 -0.0126 0.0069 1.0000
12.000 1.1339 0.05264 0.04754 -0.0128 0.0068 1.0000
12.250 1.1266 0.05675 0.05187 -0.0136 0.0067 1.0000
12.500 1.1174 0.06138 0.05672 -0.0151 0.0067 1.0000
12.750 1.1062 0.06656 0.06211 -0.0173 0.0066 1.0000
13.000 1.0932 0.07233 0.06809 -0.0201 0.0066 1.0000
13.250 1.0783 0.07875 0.07471 -0.0235 0.0066 1.0000
13.500 1.0615 0.08587 0.08204 -0.0275 0.0066 1.0000
13.750 1.0430 0.09371 0.09005 -0.0321 0.0066 1.0000
14.000 1.0230 0.10231 0.09883 -0.0372 0.0067 1.0000
14.250 1.0017 0.11190 0.10858 -0.0429 0.0068 1.0000
14.500 0.9787 0.12261 0.11940 -0.0492 0.0070 1.0000
14.750 0.9537 0.13459 0.13150 -0.0559 0.0072 1.0000
15.000 0.9255 0.14844 0.14542 -0.0632 0.0075 1.0000
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Polar data table (+)
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