GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 113 AIRFOIL (goe113-il) Reynolds number: 1,000,000 Max Cl/Cd: 97.12 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe113-il-1000000.txt Download as CSV file: xf-goe113-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 113 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4568 0.09343 0.09187 -0.0069 1.0000 0.0076
-7.750 -0.4531 0.08987 0.08833 -0.0090 1.0000 0.0076
-7.500 -0.4498 0.08625 0.08473 -0.0115 1.0000 0.0076
-7.250 -0.4412 0.08200 0.08049 -0.0160 1.0000 0.0077
-7.000 -0.4311 0.07762 0.07611 -0.0203 1.0000 0.0077
-6.750 -0.4205 0.07321 0.07169 -0.0242 1.0000 0.0077
-6.500 -0.3760 0.05338 0.05200 -0.0261 1.0000 0.0081
-6.250 -0.3723 0.05011 0.04873 -0.0268 1.0000 0.0083
-6.000 -0.3673 0.04711 0.04572 -0.0272 1.0000 0.0086
-5.750 -0.3463 0.04260 0.04114 -0.0319 0.9986 0.0093
-5.500 -0.3129 0.03710 0.03552 -0.0387 0.9966 0.0116
-5.250 -0.2819 0.03216 0.03044 -0.0434 0.9949 0.0118
-4.000 -0.1657 0.02301 0.01989 -0.0549 0.9880 0.0102
-3.750 -0.1289 0.01534 0.01139 -0.0558 0.9857 0.0112
-3.500 -0.0942 0.01122 0.00666 -0.0570 0.9837 0.0126
-3.250 -0.0620 0.01161 0.00710 -0.0581 0.9772 0.0136
-3.000 -0.0282 0.01155 0.00700 -0.0594 0.9689 0.0155
-2.750 0.0015 0.01094 0.00625 -0.0595 0.9545 0.0172
-2.500 0.0289 0.00983 0.00495 -0.0592 0.9330 0.0190
-2.250 0.0544 0.00952 0.00455 -0.0585 0.8971 0.0208
-2.000 0.0788 0.00957 0.00441 -0.0576 0.8446 0.0230
-1.750 0.1041 0.00937 0.00395 -0.0568 0.7945 0.0246
-1.500 0.1303 0.00925 0.00359 -0.0563 0.7509 0.0259
-1.250 0.1569 0.00891 0.00307 -0.0561 0.7177 0.0284
-1.000 0.1837 0.00842 0.00251 -0.0559 0.6942 0.0311
-0.750 0.2111 0.00820 0.00219 -0.0559 0.6716 0.0320
-0.500 0.2386 0.00803 0.00192 -0.0558 0.6508 0.0323
-0.250 0.2664 0.00788 0.00167 -0.0557 0.6319 0.0323
0.000 0.2940 0.00779 0.00148 -0.0557 0.6116 0.0330
0.250 0.3218 0.00771 0.00130 -0.0557 0.5913 0.0328
0.500 0.3497 0.00766 0.00116 -0.0557 0.5741 0.0332
0.750 0.3776 0.00762 0.00105 -0.0557 0.5582 0.0350
1.250 0.4333 0.00764 0.00094 -0.0557 0.5270 0.0389
1.500 0.4612 0.00767 0.00092 -0.0557 0.5137 0.0415
1.750 0.4891 0.00771 0.00094 -0.0558 0.4999 0.0477
2.000 0.5111 0.00554 0.00111 -0.0552 0.4868 1.0000
2.250 0.5389 0.00565 0.00114 -0.0552 0.4708 1.0000
2.500 0.5662 0.00583 0.00120 -0.0552 0.4418 1.0000
2.750 0.5921 0.00626 0.00130 -0.0550 0.3679 1.0000
3.000 0.6173 0.00684 0.00151 -0.0548 0.2828 1.0000
3.250 0.6435 0.00725 0.00171 -0.0547 0.2368 1.0000
3.500 0.6701 0.00757 0.00188 -0.0546 0.2065 1.0000
3.750 0.6969 0.00786 0.00204 -0.0546 0.1783 1.0000
4.000 0.7230 0.00825 0.00223 -0.0545 0.1390 1.0000
4.250 0.7489 0.00870 0.00250 -0.0543 0.1040 1.0000
4.500 0.7723 0.00963 0.00304 -0.0538 0.0174 1.0000
4.750 0.7991 0.00990 0.00335 -0.0536 0.0147 1.0000
5.000 0.8256 0.01022 0.00373 -0.0534 0.0131 1.0000
5.250 0.8516 0.01067 0.00426 -0.0531 0.0112 1.0000
5.500 0.8752 0.01162 0.00540 -0.0523 0.0096 1.0000
5.750 0.9017 0.01188 0.00567 -0.0521 0.0091 1.0000
6.000 0.9269 0.01241 0.00627 -0.0517 0.0087 1.0000
6.250 0.9515 0.01302 0.00694 -0.0512 0.0082 1.0000
6.500 0.9756 0.01369 0.00769 -0.0506 0.0076 1.0000
6.750 0.9988 0.01448 0.00856 -0.0499 0.0069 1.0000
7.000 1.0209 0.01544 0.00960 -0.0490 0.0065 1.0000
7.250 1.0414 0.01670 0.01095 -0.0478 0.0064 1.0000
7.500 1.0603 0.01861 0.01301 -0.0460 0.0070 1.0000
7.750 1.0817 0.01989 0.01436 -0.0450 0.0067 1.0000
8.000 1.1029 0.02089 0.01539 -0.0444 0.0061 1.0000
8.750 1.1592 0.02828 0.02354 -0.0395 0.0071 1.0000
9.000 1.1777 0.02951 0.02483 -0.0387 0.0065 1.0000
16.750 0.7312 0.17989 0.17848 -0.0627 0.0057 1.0000
17.000 0.7225 0.18583 0.18439 -0.0646 0.0054 1.0000
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Polar data table (+)
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