GOE 416A AIRFOIL (goe416a-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 416A AIRFOIL (goe416a-il) Reynolds number: 500,000 Max Cl/Cd: 71 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe416a-il-500000.txt Download as CSV file: xf-goe416a-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.4675 0.13663 0.13446 -0.0010 1.0000 0.0131
-13.250 -0.4766 0.13206 0.12990 -0.0017 1.0000 0.0136
-9.250 -0.7442 0.05164 0.04832 -0.0387 1.0000 0.0137
-9.000 -0.7354 0.04815 0.04478 -0.0390 1.0000 0.0140
-8.750 -0.7263 0.04514 0.04167 -0.0393 1.0000 0.0143
-8.500 -0.7156 0.04234 0.03872 -0.0394 1.0000 0.0147
-8.250 -0.7026 0.03969 0.03589 -0.0392 1.0000 0.0153
-8.000 -0.6876 0.03717 0.03315 -0.0389 1.0000 0.0161
-7.750 -0.6700 0.03509 0.03081 -0.0382 1.0000 0.0172
-7.500 -0.6460 0.03611 0.03153 -0.0371 1.0000 0.0187
-7.250 -0.6373 0.03060 0.02552 -0.0361 1.0000 0.0196
-7.000 -0.6194 0.02764 0.02257 -0.0357 1.0000 0.0208
-6.750 -0.6004 0.02597 0.02084 -0.0348 1.0000 0.0219
-6.500 -0.5814 0.02458 0.01932 -0.0335 1.0000 0.0233
-6.250 -0.5622 0.02371 0.01828 -0.0320 1.0000 0.0251
-6.000 -0.5425 0.02554 0.01991 -0.0300 1.0000 0.0269
-5.250 -0.4477 0.01665 0.01054 -0.0316 0.9950 0.0196
-5.000 -0.4127 0.01534 0.00925 -0.0330 0.9919 0.0183
-4.750 -0.3769 0.01434 0.00821 -0.0347 0.9882 0.0186
-4.500 -0.3410 0.01303 0.00687 -0.0367 0.9850 0.0195
-4.250 -0.3056 0.01203 0.00582 -0.0386 0.9803 0.0196
-4.000 -0.2718 0.01107 0.00476 -0.0401 0.9736 0.0204
-3.750 -0.2378 0.01044 0.00406 -0.0415 0.9672 0.0218
-3.500 -0.2068 0.01002 0.00357 -0.0422 0.9580 0.0240
-3.250 -0.1775 0.00973 0.00320 -0.0424 0.9473 0.0266
-3.000 -0.1495 0.00946 0.00283 -0.0421 0.9352 0.0323
-2.750 -0.1272 0.00727 0.00209 -0.0424 0.9220 0.4184
-2.500 -0.1008 0.00737 0.00227 -0.0418 0.9078 0.4803
-2.250 -0.0741 0.00745 0.00230 -0.0412 0.8925 0.4955
-2.000 -0.0471 0.00757 0.00230 -0.0407 0.8758 0.5065
-1.750 -0.0203 0.00768 0.00238 -0.0402 0.8579 0.5196
-1.500 0.0067 0.00773 0.00239 -0.0398 0.8395 0.5278
-1.250 0.0344 0.00784 0.00239 -0.0395 0.8192 0.5360
-1.000 0.0618 0.00782 0.00231 -0.0393 0.7969 0.5396
-0.750 0.0895 0.00786 0.00227 -0.0391 0.7733 0.5435
-0.500 0.1174 0.00793 0.00222 -0.0390 0.7506 0.5472
-0.250 0.1456 0.00800 0.00217 -0.0390 0.7280 0.5500
0.000 0.1735 0.00801 0.00210 -0.0389 0.7055 0.5522
0.250 0.2014 0.00807 0.00208 -0.0389 0.6832 0.5546
0.500 0.2295 0.00814 0.00208 -0.0389 0.6641 0.5571
0.750 0.2578 0.00820 0.00208 -0.0390 0.6459 0.5596
1.000 0.2861 0.00828 0.00209 -0.0391 0.6284 0.5622
1.250 0.3145 0.00837 0.00211 -0.0392 0.6111 0.5648
1.500 0.3425 0.00841 0.00213 -0.0392 0.5934 0.5674
1.750 0.3708 0.00846 0.00217 -0.0393 0.5760 0.5699
2.000 0.3991 0.00853 0.00222 -0.0394 0.5566 0.5727
2.250 0.4272 0.00863 0.00226 -0.0395 0.5353 0.5757
2.500 0.4555 0.00874 0.00231 -0.0396 0.5133 0.5788
2.750 0.4832 0.00883 0.00237 -0.0396 0.4920 0.5820
3.000 0.5111 0.00894 0.00248 -0.0397 0.4722 0.5851
3.250 0.5390 0.00908 0.00259 -0.0397 0.4548 0.5886
3.500 0.5670 0.00922 0.00271 -0.0398 0.4402 0.5923
3.750 0.5947 0.00934 0.00284 -0.0399 0.4247 0.5960
4.000 0.6220 0.00952 0.00300 -0.0399 0.4006 0.6003
4.250 0.6494 0.00972 0.00315 -0.0399 0.3744 0.6052
4.500 0.6769 0.00989 0.00330 -0.0399 0.3451 0.6100
4.750 0.7037 0.01014 0.00348 -0.0399 0.3060 0.6153
5.000 0.7300 0.01050 0.00372 -0.0398 0.2691 0.6214
5.250 0.7562 0.01084 0.00402 -0.0397 0.2455 0.6279
5.500 0.7825 0.01118 0.00433 -0.0396 0.2248 0.6361
5.750 0.8087 0.01149 0.00465 -0.0395 0.2039 0.6459
6.000 0.8350 0.01176 0.00494 -0.0395 0.1792 0.6588
6.250 0.8598 0.01225 0.00529 -0.0393 0.1277 0.6771
6.500 0.8833 0.01284 0.00583 -0.0389 0.0937 0.7078
6.750 0.9046 0.01312 0.00638 -0.0378 0.0674 0.8118
7.000 0.9267 0.01347 0.00689 -0.0368 0.0498 1.0000
7.250 0.9509 0.01416 0.00754 -0.0364 0.0368 1.0000
7.500 0.9742 0.01496 0.00833 -0.0358 0.0268 1.0000
7.750 0.9977 0.01569 0.00907 -0.0353 0.0218 1.0000
8.000 1.0201 0.01655 0.01001 -0.0346 0.0190 1.0000
8.250 1.0436 0.01718 0.01070 -0.0341 0.0170 1.0000
8.500 1.0639 0.01822 0.01179 -0.0332 0.0153 1.0000
8.750 1.0809 0.01964 0.01334 -0.0318 0.0143 1.0000
9.000 1.1014 0.02055 0.01436 -0.0309 0.0136 1.0000
9.250 1.1203 0.02162 0.01553 -0.0298 0.0129 1.0000
9.500 1.1385 0.02272 0.01672 -0.0287 0.0123 1.0000
9.750 1.1565 0.02375 0.01783 -0.0276 0.0116 1.0000
10.000 1.1714 0.02510 0.01924 -0.0263 0.0109 1.0000
10.250 1.1753 0.02811 0.02247 -0.0237 0.0103 1.0000
10.500 1.1893 0.02949 0.02401 -0.0222 0.0100 1.0000
10.750 1.2013 0.03099 0.02568 -0.0206 0.0098 1.0000
11.000 1.2087 0.03266 0.02752 -0.0184 0.0096 1.0000
11.250 1.2127 0.03456 0.02961 -0.0160 0.0094 1.0000
11.500 1.2142 0.03672 0.03196 -0.0139 0.0092 1.0000
11.750 1.2133 0.03917 0.03461 -0.0120 0.0091 1.0000
12.000 1.2100 0.04195 0.03760 -0.0106 0.0090 1.0000
12.250 1.2045 0.04512 0.04098 -0.0098 0.0088 1.0000
12.500 1.1959 0.04890 0.04497 -0.0097 0.0088 1.0000
12.750 1.1851 0.05322 0.04951 -0.0104 0.0087 1.0000
13.000 1.1717 0.05826 0.05477 -0.0120 0.0087 1.0000
13.250 1.1560 0.06402 0.06075 -0.0145 0.0087 1.0000
13.500 1.1381 0.07051 0.06744 -0.0179 0.0088 1.0000
13.750 1.1176 0.07789 0.07502 -0.0220 0.0088 1.0000
14.000 1.0957 0.08596 0.08327 -0.0268 0.0089 1.0000
14.250 1.0723 0.09485 0.09233 -0.0322 0.0090 1.0000
14.500 1.0479 0.10451 0.10214 -0.0381 0.0091 1.0000
14.750 1.0221 0.11533 0.11311 -0.0447 0.0093 1.0000
15.000 0.9935 0.12761 0.12551 -0.0519 0.0095 1.0000
15.250 0.9572 0.14282 0.14083 -0.0604 0.0098 1.0000
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Polar data table (+)
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