GOE 411 AIRFOIL (goe411-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 411 AIRFOIL (goe411-il) Reynolds number: 50,000 Max Cl/Cd: 27.98 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe411-il-50000-n5.txt Download as CSV file: xf-goe411-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 411 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
------ -------- --------- --------- -------- -------- --------
-12.500 -0.6461 0.09538 0.08815 -0.0423 1.0000 0.0301
-12.250 -0.6682 0.08826 0.08092 -0.0464 1.0000 0.0297
-12.000 -0.6895 0.08221 0.07472 -0.0496 1.0000 0.0294
-11.750 -0.7096 0.07703 0.06939 -0.0516 1.0000 0.0291
-11.500 -0.7291 0.07255 0.06469 -0.0525 1.0000 0.0288
-11.250 -0.7464 0.06872 0.06066 -0.0523 1.0000 0.0287
-11.000 -0.7623 0.06537 0.05706 -0.0511 1.0000 0.0285
-10.750 -0.7754 0.06246 0.05391 -0.0491 1.0000 0.0283
-10.500 -0.7845 0.05974 0.05096 -0.0465 1.0000 0.0282
-10.250 -0.7908 0.05713 0.04809 -0.0436 1.0000 0.0281
-10.000 -0.7932 0.05467 0.04537 -0.0406 1.0000 0.0281
-9.750 -0.7908 0.05226 0.04272 -0.0379 1.0000 0.0282
-9.500 -0.7822 0.04977 0.03997 -0.0357 1.0000 0.0284
-9.250 -0.7690 0.04737 0.03738 -0.0341 1.0000 0.0290
-9.000 -0.7531 0.04516 0.03499 -0.0327 1.0000 0.0302
-8.750 -0.7354 0.04323 0.03287 -0.0313 1.0000 0.0315
-8.500 -0.7176 0.04143 0.03077 -0.0297 1.0000 0.0330
-8.250 -0.7048 0.03992 0.02909 -0.0273 1.0000 0.0336
-8.000 -0.6976 0.03852 0.02753 -0.0242 1.0000 0.0343
-7.750 -0.6940 0.03725 0.02612 -0.0206 1.0000 0.0349
-7.500 -0.6908 0.03597 0.02467 -0.0169 1.0000 0.0355
-7.250 -0.6900 0.03455 0.02316 -0.0130 1.0000 0.0366
-7.000 -0.6890 0.03312 0.02164 -0.0091 1.0000 0.0386
-6.750 -0.6846 0.03189 0.02029 -0.0056 1.0000 0.0411
-6.500 -0.6778 0.03076 0.01899 -0.0022 1.0000 0.0444
-6.250 -0.6708 0.02952 0.01769 0.0011 1.0000 0.0481
-6.000 -0.6617 0.02840 0.01651 0.0041 1.0000 0.0539
-5.750 -0.6573 0.02666 0.01517 0.0075 1.0000 0.0900
-5.500 -0.6581 0.02476 0.01429 0.0115 1.0000 0.2435
-5.250 -0.6527 0.02381 0.01380 0.0152 1.0000 0.3251
-5.000 -0.6443 0.02322 0.01338 0.0188 1.0000 0.3868
-4.750 -0.6357 0.02280 0.01312 0.0224 1.0000 0.4441
-4.500 -0.6293 0.02249 0.01299 0.0267 1.0000 0.5040
-4.250 -0.6184 0.02225 0.01282 0.0301 1.0000 0.5488
-4.000 -0.6070 0.02209 0.01281 0.0339 1.0000 0.5985
-3.750 -0.5879 0.02193 0.01259 0.0358 1.0000 0.6274
-3.500 -0.5673 0.02179 0.01237 0.0373 1.0000 0.6518
-3.250 -0.5466 0.02172 0.01217 0.0389 1.0000 0.6772
-3.000 -0.5241 0.02174 0.01215 0.0402 1.0000 0.7031
-2.750 -0.5015 0.02182 0.01218 0.0414 1.0000 0.7312
-2.500 -0.4774 0.02198 0.01225 0.0424 1.0000 0.7607
-2.250 -0.4503 0.02222 0.01244 0.0428 1.0000 0.7895
-2.000 -0.4189 0.02255 0.01268 0.0423 1.0000 0.8157
-1.750 -0.3875 0.02285 0.01288 0.0415 1.0000 0.8395
-1.500 -0.3345 0.02337 0.01322 0.0363 0.9946 0.8578
-1.250 -0.2778 0.02385 0.01354 0.0303 0.9869 0.8743
-1.000 -0.2203 0.02424 0.01378 0.0240 0.9790 0.8887
-0.750 -0.1637 0.02455 0.01397 0.0177 0.9702 0.9018
-0.500 -0.1084 0.02476 0.01411 0.0117 0.9605 0.9146
-0.250 -0.0541 0.02489 0.01418 0.0059 0.9501 0.9270
0.000 -0.0002 0.02493 0.01421 0.0000 0.9390 0.9389
0.250 0.0540 0.02488 0.01418 -0.0058 0.9271 0.9501
0.500 0.1083 0.02476 0.01411 -0.0117 0.9147 0.9605
0.750 0.1636 0.02454 0.01398 -0.0177 0.9018 0.9702
1.000 0.2203 0.02423 0.01377 -0.0240 0.8887 0.9789
1.250 0.2779 0.02384 0.01353 -0.0303 0.8743 0.9870
1.500 0.3345 0.02336 0.01321 -0.0363 0.8578 0.9946
1.750 0.3874 0.02285 0.01288 -0.0415 0.8395 1.0000
2.000 0.4189 0.02254 0.01268 -0.0423 0.8157 1.0000
2.250 0.4502 0.02222 0.01243 -0.0428 0.7896 1.0000
2.500 0.4774 0.02197 0.01228 -0.0424 0.7607 1.0000
2.750 0.5013 0.02182 0.01217 -0.0414 0.7312 1.0000
3.000 0.5241 0.02174 0.01214 -0.0402 0.7032 1.0000
3.250 0.5465 0.02171 0.01217 -0.0389 0.6771 1.0000
3.500 0.5671 0.02179 0.01235 -0.0373 0.6518 1.0000
3.750 0.5878 0.02192 0.01258 -0.0358 0.6274 1.0000
4.000 0.6067 0.02208 0.01280 -0.0338 0.5979 1.0000
4.250 0.6183 0.02224 0.01281 -0.0301 0.5488 1.0000
4.500 0.6289 0.02248 0.01297 -0.0266 0.5030 1.0000
4.750 0.6352 0.02280 0.01309 -0.0223 0.4428 1.0000
5.000 0.6443 0.02322 0.01337 -0.0188 0.3869 1.0000
5.250 0.6522 0.02382 0.01380 -0.0151 0.3221 1.0000
5.500 0.6579 0.02477 0.01429 -0.0114 0.2425 1.0000
6.000 0.6618 0.02838 0.01649 -0.0041 0.0540 1.0000
6.250 0.6707 0.02953 0.01769 -0.0011 0.0480 1.0000
6.500 0.6779 0.03074 0.01897 0.0022 0.0446 1.0000
6.750 0.6850 0.03184 0.02024 0.0055 0.0417 1.0000
7.000 0.6892 0.03310 0.02162 0.0091 0.0388 1.0000
7.250 0.6907 0.03449 0.02310 0.0129 0.0368 1.0000
7.500 0.6910 0.03595 0.02465 0.0169 0.0355 1.0000
7.750 0.6944 0.03721 0.02608 0.0205 0.0349 1.0000
8.000 0.6979 0.03851 0.02752 0.0241 0.0343 1.0000
8.250 0.7049 0.03988 0.02905 0.0273 0.0337 1.0000
8.500 0.7178 0.04142 0.03076 0.0296 0.0329 1.0000
8.750 0.7354 0.04317 0.03281 0.0313 0.0317 1.0000
9.000 0.7528 0.04510 0.03492 0.0327 0.0303 1.0000
9.250 0.7688 0.04735 0.03733 0.0341 0.0289 1.0000
9.500 0.7825 0.04982 0.04002 0.0356 0.0283 1.0000
9.750 0.7911 0.05224 0.04270 0.0378 0.0282 1.0000
10.000 0.7935 0.05466 0.04536 0.0405 0.0281 1.0000
10.250 0.7912 0.05720 0.04816 0.0435 0.0281 1.0000
10.500 0.7851 0.05978 0.05099 0.0464 0.0282 1.0000
10.750 0.7761 0.06242 0.05387 0.0490 0.0283 1.0000
11.000 0.7626 0.06543 0.05713 0.0510 0.0284 1.0000
11.250 0.7471 0.06875 0.06068 0.0523 0.0286 1.0000
11.500 0.7295 0.07259 0.06474 0.0525 0.0288 1.0000
11.750 0.7100 0.07711 0.06946 0.0515 0.0291 1.0000
12.000 0.6908 0.08212 0.07463 0.0495 0.0294 1.0000
12.250 0.6687 0.08835 0.08100 0.0463 0.0297 1.0000
12.500 0.6472 0.09537 0.08813 0.0421 0.0301 1.0000
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Polar data table (+)
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