GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 611 AIRFOIL (goe611-il) Reynolds number: 500,000 Max Cl/Cd: 97.01 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe611-il-500000-n5.txt Download as CSV file: xf-goe611-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 611 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1358 0.08925 0.08644 -0.0736 0.8327 0.0082
-9.000 -0.1323 0.08546 0.08262 -0.0756 0.8247 0.0083
-8.750 -0.1295 0.08157 0.07870 -0.0778 0.8172 0.0084
-8.500 -0.1228 0.07882 0.07593 -0.0795 0.8103 0.0085
-8.250 -0.1143 0.07650 0.07359 -0.0811 0.8035 0.0087
-8.000 -0.1081 0.07361 0.07068 -0.0834 0.7972 0.0089
-7.750 -0.1018 0.07048 0.06754 -0.0866 0.7907 0.0092
-7.500 -0.0918 0.06654 0.06355 -0.0913 0.7849 0.0094
-7.250 -0.0801 0.06189 0.05886 -0.0967 0.7791 0.0096
-7.000 -0.0667 0.05682 0.05371 -0.1021 0.7732 0.0096
-6.750 -0.0513 0.05159 0.04835 -0.1068 0.7680 0.0098
-6.500 -0.0344 0.04585 0.04244 -0.1109 0.7624 0.0101
-6.250 -0.0215 0.03499 0.03108 -0.1150 0.7575 0.0111
-6.000 0.0019 0.03365 0.02962 -0.1155 0.7520 0.0114
-5.500 0.0424 0.02218 0.01707 -0.1156 0.7411 0.0143
-5.250 0.0687 0.02172 0.01651 -0.1158 0.7350 0.0146
-5.000 0.0953 0.02127 0.01596 -0.1159 0.7280 0.0150
-4.750 0.1218 0.02040 0.01492 -0.1160 0.7214 0.0160
-4.500 0.1472 0.01694 0.01086 -0.1154 0.7146 0.0185
-4.250 0.1744 0.01668 0.01050 -0.1155 0.7071 0.0190
-4.000 0.2020 0.01637 0.01011 -0.1156 0.6987 0.0199
-3.750 0.2294 0.01561 0.00912 -0.1155 0.6907 0.0215
-3.500 0.2573 0.01485 0.00807 -0.1154 0.6815 0.0227
-3.250 0.2843 0.01411 0.00722 -0.1154 0.6724 0.0237
-3.000 0.3115 0.01383 0.00686 -0.1154 0.6625 0.0245
-2.750 0.3391 0.01344 0.00639 -0.1155 0.6521 0.0253
-2.500 0.3665 0.01303 0.00587 -0.1154 0.6418 0.0262
-2.000 0.4211 0.01250 0.00510 -0.1152 0.6184 0.0281
-1.750 0.4479 0.01228 0.00476 -0.1150 0.6033 0.0286
-1.500 0.4743 0.01212 0.00448 -0.1148 0.5864 0.0290
-1.250 0.5006 0.01195 0.00419 -0.1146 0.5703 0.0293
-1.000 0.5269 0.01154 0.00371 -0.1144 0.5576 0.0301
-0.750 0.5538 0.01134 0.00347 -0.1144 0.5476 0.0312
-0.500 0.5807 0.01123 0.00330 -0.1143 0.5380 0.0317
-0.250 0.6080 0.01113 0.00316 -0.1143 0.5287 0.0323
0.000 0.6351 0.01108 0.00305 -0.1143 0.5201 0.0331
0.250 0.6623 0.01105 0.00297 -0.1143 0.5111 0.0338
0.500 0.6895 0.01105 0.00292 -0.1143 0.5032 0.0346
0.750 0.7165 0.01106 0.00289 -0.1143 0.4945 0.0353
1.000 0.7437 0.01109 0.00287 -0.1143 0.4866 0.0363
1.250 0.7705 0.01116 0.00288 -0.1142 0.4781 0.0375
1.500 0.7976 0.01121 0.00289 -0.1142 0.4699 0.0390
1.750 0.8241 0.01129 0.00292 -0.1141 0.4612 0.0416
2.000 0.8510 0.01136 0.00298 -0.1141 0.4527 0.0471
2.250 0.8732 0.00950 0.00330 -0.1135 0.4445 1.0000
2.500 0.8996 0.00965 0.00338 -0.1134 0.4357 1.0000
2.750 0.9253 0.00983 0.00348 -0.1131 0.4263 1.0000
3.000 0.9504 0.01004 0.00360 -0.1128 0.4145 1.0000
3.250 0.9760 0.01023 0.00373 -0.1126 0.4037 1.0000
3.500 1.0013 0.01043 0.00387 -0.1123 0.3938 1.0000
3.750 1.0260 0.01066 0.00403 -0.1119 0.3836 1.0000
4.000 1.0512 0.01086 0.00419 -0.1117 0.3737 1.0000
4.250 1.0758 0.01109 0.00437 -0.1113 0.3638 1.0000
4.500 1.0998 0.01135 0.00457 -0.1108 0.3531 1.0000
4.750 1.1237 0.01161 0.00479 -0.1103 0.3414 1.0000
5.000 1.1470 0.01190 0.00502 -0.1098 0.3292 1.0000
5.250 1.1697 0.01221 0.00527 -0.1091 0.3167 1.0000
5.500 1.1919 0.01255 0.00555 -0.1083 0.3039 1.0000
5.750 1.2134 0.01290 0.00585 -0.1075 0.2910 1.0000
6.000 1.2334 0.01333 0.00619 -0.1064 0.2735 1.0000
6.250 1.2530 0.01375 0.00655 -0.1052 0.2585 1.0000
6.500 1.2713 0.01422 0.00696 -0.1038 0.2417 1.0000
6.750 1.2834 0.01491 0.00750 -0.1013 0.2151 1.0000
7.000 1.2823 0.01627 0.00850 -0.0967 0.1557 1.0000
7.250 1.2896 0.01729 0.00935 -0.0936 0.1309 1.0000
7.500 1.3038 0.01796 0.00999 -0.0918 0.1231 1.0000
7.750 1.3203 0.01851 0.01057 -0.0904 0.1186 1.0000
8.000 1.3367 0.01909 0.01116 -0.0890 0.1152 1.0000
8.250 1.3518 0.01975 0.01184 -0.0875 0.1116 1.0000
8.500 1.3658 0.02050 0.01260 -0.0859 0.1073 1.0000
8.750 1.3822 0.02112 0.01327 -0.0847 0.1041 1.0000
9.000 1.3985 0.02176 0.01395 -0.0835 0.1000 1.0000
9.250 1.4100 0.02274 0.01490 -0.0819 0.0925 1.0000
9.500 1.4098 0.02456 0.01648 -0.0790 0.0559 1.0000
9.750 1.3965 0.02750 0.01929 -0.0751 0.0172 1.0000
10.000 1.4052 0.02889 0.02074 -0.0736 0.0142 1.0000
10.250 1.4156 0.03018 0.02212 -0.0723 0.0128 1.0000
10.500 1.4254 0.03156 0.02358 -0.0712 0.0117 1.0000
10.750 1.4333 0.03312 0.02523 -0.0699 0.0107 1.0000
11.000 1.4386 0.03494 0.02714 -0.0685 0.0098 1.0000
11.250 1.4467 0.03655 0.02884 -0.0675 0.0094 1.0000
11.500 1.4534 0.03832 0.03070 -0.0664 0.0089 1.0000
11.750 1.4589 0.04025 0.03273 -0.0654 0.0084 1.0000
12.000 1.4634 0.04231 0.03488 -0.0644 0.0080 1.0000
12.250 1.4663 0.04460 0.03726 -0.0634 0.0076 1.0000
12.500 1.4659 0.04730 0.04007 -0.0625 0.0072 1.0000
12.750 1.4687 0.04976 0.04262 -0.0618 0.0069 1.0000
13.000 1.4709 0.05232 0.04528 -0.0612 0.0066 1.0000
13.250 1.4716 0.05513 0.04818 -0.0607 0.0064 1.0000
13.500 1.4714 0.05812 0.05129 -0.0603 0.0061 1.0000
13.750 1.4703 0.06128 0.05455 -0.0600 0.0060 1.0000
14.000 1.4687 0.06457 0.05793 -0.0598 0.0058 1.0000
14.250 1.4660 0.06805 0.06151 -0.0598 0.0056 1.0000
14.500 1.4622 0.07175 0.06532 -0.0598 0.0055 1.0000
14.750 1.4569 0.07572 0.06939 -0.0600 0.0054 1.0000
15.000 1.4495 0.08004 0.07383 -0.0603 0.0052 1.0000
15.250 1.4390 0.08483 0.07873 -0.0607 0.0051 1.0000
15.500 1.4320 0.08922 0.08324 -0.0612 0.0051 1.0000
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Polar data table (+)
Polar graphs
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