GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 590 AIRFOIL (goe590-il) Reynolds number: 1,000,000 Max Cl/Cd: 114.64 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe590-il-1000000.txt Download as CSV file: xf-goe590-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 590 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3617 0.10815 0.10664 -0.0060 1.0000 0.0061
-9.750 -0.3579 0.10453 0.10303 -0.0072 1.0000 0.0061
-9.250 -0.3552 0.09549 0.09401 -0.0091 1.0000 0.0064
-9.000 -0.3507 0.09213 0.09066 -0.0098 1.0000 0.0066
-8.750 -0.3464 0.08893 0.08747 -0.0106 1.0000 0.0068
-8.500 -0.3432 0.08547 0.08401 -0.0115 1.0000 0.0069
-8.250 -0.3403 0.08207 0.08062 -0.0125 1.0000 0.0071
-8.000 -0.3379 0.07863 0.07720 -0.0134 1.0000 0.0072
-7.750 -0.3362 0.07528 0.07386 -0.0143 1.0000 0.0073
-7.500 -0.3354 0.07195 0.07056 -0.0152 1.0000 0.0075
-7.250 -0.3363 0.06872 0.06734 -0.0160 1.0000 0.0075
-7.000 -0.3403 0.06574 0.06439 -0.0166 1.0000 0.0077
-6.750 -0.3399 0.06230 0.06096 -0.0182 1.0000 0.0078
-6.500 -0.3224 0.05710 0.05573 -0.0242 0.9679 0.0082
-6.250 -0.2983 0.05145 0.04993 -0.0311 0.9082 0.0088
-6.000 -0.2906 0.04802 0.04629 -0.0322 0.8583 0.0092
-4.500 -0.1874 0.04120 0.03838 -0.0467 0.7553 0.0101
-4.250 -0.1663 0.03739 0.03437 -0.0467 0.7394 0.0101
-4.000 -0.1501 0.03205 0.02877 -0.0466 0.7268 0.0114
-3.750 -0.1272 0.03098 0.02758 -0.0466 0.7128 0.0132
-3.000 -0.0667 0.01054 0.00496 -0.0391 0.6847 0.0111
-2.750 -0.0399 0.01052 0.00488 -0.0389 0.6702 0.0127
-2.500 -0.0136 0.01018 0.00441 -0.0385 0.6550 0.0141
-2.250 0.0128 0.00988 0.00398 -0.0381 0.6379 0.0155
-2.000 0.0381 0.00922 0.00316 -0.0376 0.6184 0.0181
-1.750 0.0648 0.00930 0.00317 -0.0374 0.5979 0.0207
-1.500 0.0911 0.00916 0.00292 -0.0370 0.5799 0.0232
-1.250 0.1180 0.00923 0.00291 -0.0368 0.5655 0.0249
-1.000 0.1427 0.00862 0.00216 -0.0362 0.5546 0.0282
-0.750 0.1682 0.00834 0.00182 -0.0357 0.5451 0.0303
-0.500 0.1943 0.00815 0.00158 -0.0353 0.5370 0.0329
-0.250 0.2205 0.00806 0.00144 -0.0350 0.5297 0.0358
0.000 0.2470 0.00796 0.00130 -0.0348 0.5234 0.0370
0.250 0.2734 0.00789 0.00117 -0.0345 0.5176 0.0375
0.500 0.3000 0.00781 0.00107 -0.0342 0.5126 0.0378
0.750 0.3266 0.00774 0.00096 -0.0340 0.5074 0.0384
1.000 0.3531 0.00772 0.00088 -0.0337 0.5024 0.0411
1.250 0.3800 0.00764 0.00084 -0.0335 0.4981 0.0507
1.500 0.4054 0.00741 0.00091 -0.0331 0.4938 0.1656
1.750 0.4126 0.00573 0.00097 -0.0292 0.4904 0.8456
2.000 0.4849 0.00566 0.00120 -0.0391 0.4830 0.9857
2.250 0.5304 0.00580 0.00127 -0.0432 0.4742 0.9937
2.500 0.5774 0.00583 0.00131 -0.0477 0.4683 0.9979
2.750 0.6154 0.00590 0.00135 -0.0501 0.4619 1.0000
3.000 0.6405 0.00593 0.00139 -0.0496 0.4537 1.0000
3.250 0.6652 0.00601 0.00142 -0.0490 0.4382 1.0000
3.500 0.6898 0.00610 0.00149 -0.0485 0.4220 1.0000
3.750 0.7142 0.00623 0.00157 -0.0478 0.3994 1.0000
4.000 0.7360 0.00666 0.00171 -0.0469 0.3296 1.0000
4.250 0.7440 0.00894 0.00273 -0.0441 0.0447 1.0000
4.500 0.7665 0.00933 0.00306 -0.0431 0.0269 1.0000
4.750 0.7897 0.00962 0.00338 -0.0423 0.0207 1.0000
5.000 0.8110 0.01017 0.00402 -0.0410 0.0147 1.0000
5.250 0.8341 0.01045 0.00432 -0.0402 0.0133 1.0000
5.500 0.8562 0.01085 0.00478 -0.0392 0.0118 1.0000
5.750 0.8777 0.01132 0.00530 -0.0380 0.0105 1.0000
6.000 0.8924 0.01259 0.00673 -0.0356 0.0090 1.0000
6.250 0.9088 0.01359 0.00785 -0.0336 0.0083 1.0000
6.500 0.9297 0.01407 0.00839 -0.0325 0.0078 1.0000
6.750 0.9471 0.01500 0.00940 -0.0307 0.0075 1.0000
7.000 0.9659 0.01578 0.01025 -0.0291 0.0069 1.0000
7.250 0.9828 0.01693 0.01149 -0.0273 0.0066 1.0000
7.500 1.0034 0.01744 0.01204 -0.0262 0.0059 1.0000
7.750 1.0220 0.01828 0.01293 -0.0249 0.0054 1.0000
8.000 1.0367 0.02023 0.01501 -0.0228 0.0051 1.0000
8.250 1.0522 0.02286 0.01788 -0.0208 0.0051 1.0000
8.500 1.0325 0.04007 0.03622 -0.0133 0.0085 1.0000
8.750 1.0389 0.04311 0.03950 -0.0109 0.0085 1.0000
9.000 1.0431 0.04600 0.04262 -0.0085 0.0085 1.0000
9.250 1.0479 0.04843 0.04528 -0.0061 0.0084 1.0000
9.500 1.0767 0.04784 0.04478 -0.0049 0.0070 1.0000
9.750 1.0772 0.05095 0.04809 -0.0025 0.0067 1.0000
10.000 1.0758 0.05354 0.05084 -0.0003 0.0062 1.0000
10.250 1.0654 0.05687 0.05434 0.0025 0.0062 1.0000
10.500 1.0540 0.05875 0.05631 0.0056 0.0059 1.0000
10.750 1.0328 0.06246 0.06017 0.0075 0.0061 1.0000
11.000 1.0197 0.06534 0.06314 0.0076 0.0059 1.0000
11.250 0.9989 0.07015 0.06808 0.0062 0.0059 1.0000
11.500 0.9762 0.07610 0.07416 0.0035 0.0061 1.0000
11.750 0.9542 0.08269 0.08086 -0.0004 0.0063 1.0000
12.000 0.9346 0.08969 0.08795 -0.0050 0.0063 1.0000
12.250 0.9117 0.09860 0.09693 -0.0109 0.0065 1.0000
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Polar data table (+)
Polar graphs
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