GOE 474 AIRFOIL (goe474-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 474 AIRFOIL (goe474-il) Reynolds number: 1,000,000 Max Cl/Cd: 115.79 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe474-il-1000000.txt Download as CSV file: xf-goe474-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 474 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3619 0.10424 0.10262 -0.0242 1.0000 0.0126
-10.500 -0.3647 0.09902 0.09740 -0.0262 1.0000 0.0141
-9.250 -0.7466 0.02663 0.02394 -0.0528 1.0000 0.0080
-9.000 -0.7209 0.02442 0.02149 -0.0544 0.9986 0.0082
-8.750 -0.6917 0.02252 0.01937 -0.0562 0.9967 0.0085
-8.500 -0.6627 0.02041 0.01697 -0.0579 0.9949 0.0087
-8.250 -0.6345 0.01873 0.01504 -0.0590 0.9928 0.0091
-8.000 -0.6058 0.01742 0.01353 -0.0599 0.9902 0.0095
-7.750 -0.5760 0.01612 0.01201 -0.0609 0.9878 0.0099
-7.500 -0.5440 0.01537 0.01111 -0.0621 0.9860 0.0103
-7.250 -0.5142 0.01362 0.00909 -0.0632 0.9843 0.0112
-7.000 -0.4817 0.01303 0.00843 -0.0644 0.9827 0.0118
-6.750 -0.4537 0.01250 0.00784 -0.0646 0.9785 0.0124
-6.500 -0.4222 0.01193 0.00719 -0.0654 0.9759 0.0131
-6.250 -0.3894 0.01146 0.00663 -0.0666 0.9739 0.0139
-6.000 -0.3568 0.01067 0.00572 -0.0677 0.9719 0.0149
-5.750 -0.3242 0.01011 0.00512 -0.0688 0.9696 0.0163
-5.500 -0.2959 0.00991 0.00491 -0.0689 0.9634 0.0176
-5.250 -0.2637 0.00966 0.00462 -0.0698 0.9592 0.0189
-5.000 -0.2334 0.00905 0.00392 -0.0703 0.9542 0.0207
-4.750 -0.2051 0.00869 0.00354 -0.0704 0.9470 0.0228
-4.500 -0.1748 0.00844 0.00325 -0.0708 0.9410 0.0248
-4.250 -0.1468 0.00833 0.00310 -0.0707 0.9326 0.0263
-4.000 -0.1199 0.00785 0.00253 -0.0704 0.9246 0.0297
-3.750 -0.0928 0.00760 0.00223 -0.0701 0.9162 0.0324
-3.500 -0.0659 0.00743 0.00201 -0.0698 0.9080 0.0348
-3.250 -0.0388 0.00726 0.00177 -0.0694 0.9003 0.0377
-3.000 -0.0125 0.00704 0.00156 -0.0690 0.8921 0.0456
-2.750 0.0133 0.00674 0.00144 -0.0685 0.8845 0.0923
-2.500 0.0401 0.00667 0.00138 -0.0682 0.8760 0.1051
-2.250 0.0670 0.00664 0.00132 -0.0678 0.8669 0.1130
-2.000 0.0936 0.00660 0.00126 -0.0674 0.8563 0.1192
-1.750 0.1205 0.00658 0.00120 -0.0670 0.8458 0.1230
-1.500 0.1472 0.00652 0.00113 -0.0667 0.8366 0.1289
-1.250 0.1739 0.00649 0.00108 -0.0663 0.8273 0.1347
-1.000 0.2006 0.00646 0.00102 -0.0659 0.8153 0.1396
-0.750 0.2271 0.00641 0.00097 -0.0654 0.8033 0.1471
-0.500 0.2538 0.00640 0.00093 -0.0650 0.7920 0.1523
-0.250 0.2802 0.00636 0.00089 -0.0646 0.7798 0.1623
0.000 0.3063 0.00633 0.00087 -0.0641 0.7650 0.1754
0.250 0.3323 0.00630 0.00084 -0.0636 0.7478 0.1915
0.500 0.3582 0.00625 0.00083 -0.0631 0.7322 0.2192
0.750 0.3831 0.00604 0.00083 -0.0625 0.7179 0.3123
1.000 0.3990 0.00501 0.00088 -0.0604 0.7022 0.7130
1.250 0.4374 0.00452 0.00103 -0.0623 0.6860 0.9674
1.500 0.4861 0.00466 0.00109 -0.0668 0.6657 0.9868
1.750 0.5358 0.00482 0.00114 -0.0716 0.6402 0.9973
2.000 0.5715 0.00497 0.00118 -0.0734 0.6101 1.0000
2.250 0.5940 0.00513 0.00122 -0.0722 0.5771 1.0000
2.500 0.6144 0.00544 0.00130 -0.0706 0.5152 1.0000
2.750 0.6317 0.00601 0.00148 -0.0686 0.4240 1.0000
3.000 0.6523 0.00639 0.00165 -0.0671 0.3737 1.0000
3.250 0.6738 0.00670 0.00179 -0.0658 0.3329 1.0000
3.500 0.6955 0.00703 0.00195 -0.0646 0.2935 1.0000
3.750 0.7166 0.00739 0.00213 -0.0633 0.2487 1.0000
4.000 0.7376 0.00779 0.00233 -0.0620 0.2058 1.0000
4.250 0.7596 0.00811 0.00253 -0.0608 0.1779 1.0000
4.500 0.7822 0.00840 0.00272 -0.0597 0.1573 1.0000
4.750 0.8047 0.00870 0.00293 -0.0587 0.1371 1.0000
5.000 0.8278 0.00896 0.00313 -0.0577 0.1210 1.0000
5.250 0.8509 0.00921 0.00333 -0.0568 0.1083 1.0000
5.500 0.8738 0.00949 0.00355 -0.0558 0.0971 1.0000
5.750 0.8973 0.00973 0.00377 -0.0550 0.0896 1.0000
6.000 0.9207 0.00997 0.00401 -0.0541 0.0826 1.0000
6.250 0.9438 0.01025 0.00426 -0.0532 0.0740 1.0000
6.500 0.9669 0.01054 0.00451 -0.0523 0.0626 1.0000
6.750 0.9878 0.01104 0.00487 -0.0511 0.0408 1.0000
7.000 1.0087 0.01158 0.00534 -0.0499 0.0272 1.0000
7.250 1.0298 0.01210 0.00584 -0.0486 0.0204 1.0000
7.500 1.0517 0.01255 0.00628 -0.0475 0.0174 1.0000
7.750 1.0728 0.01309 0.00687 -0.0463 0.0151 1.0000
8.000 1.0951 0.01349 0.00731 -0.0453 0.0139 1.0000
8.250 1.1168 0.01396 0.00781 -0.0443 0.0129 1.0000
8.500 1.1362 0.01468 0.00859 -0.0428 0.0116 1.0000
8.750 1.1551 0.01543 0.00944 -0.0413 0.0109 1.0000
9.000 1.1760 0.01593 0.01000 -0.0402 0.0104 1.0000
9.250 1.1963 0.01649 0.01063 -0.0390 0.0099 1.0000
9.500 1.2160 0.01708 0.01128 -0.0377 0.0094 1.0000
9.750 1.2360 0.01762 0.01186 -0.0365 0.0089 1.0000
10.000 1.2528 0.01845 0.01275 -0.0349 0.0084 1.0000
10.250 1.2593 0.02024 0.01472 -0.0316 0.0077 1.0000
10.500 1.2764 0.02094 0.01550 -0.0301 0.0076 1.0000
10.750 1.2940 0.02155 0.01619 -0.0286 0.0074 1.0000
11.000 1.3072 0.02251 0.01726 -0.0264 0.0072 1.0000
11.250 1.3194 0.02332 0.01815 -0.0241 0.0070 1.0000
11.500 1.3268 0.02438 0.01934 -0.0211 0.0068 1.0000
11.750 1.3352 0.02536 0.02042 -0.0184 0.0066 1.0000
12.000 1.3430 0.02637 0.02153 -0.0158 0.0065 1.0000
12.250 1.3475 0.02768 0.02296 -0.0130 0.0063 1.0000
12.500 1.3536 0.02887 0.02425 -0.0105 0.0062 1.0000
12.750 1.3576 0.03024 0.02573 -0.0082 0.0061 1.0000
13.000 1.3579 0.03200 0.02762 -0.0058 0.0060 1.0000
13.250 1.3613 0.03352 0.02922 -0.0039 0.0058 1.0000
13.500 1.3605 0.03550 0.03131 -0.0021 0.0057 1.0000
13.750 1.3545 0.03813 0.03407 -0.0004 0.0055 1.0000
14.000 1.3483 0.04099 0.03708 0.0008 0.0056 1.0000
14.250 1.3289 0.04555 0.04185 0.0015 0.0054 1.0000
14.500 1.3182 0.04948 0.04594 0.0014 0.0054 1.0000
14.750 1.3130 0.05301 0.04959 0.0008 0.0055 1.0000
15.000 1.2752 0.06156 0.05841 -0.0020 0.0053 1.0000
15.250 1.2625 0.06714 0.06414 -0.0046 0.0053 1.0000
15.500 1.2303 0.07697 0.07419 -0.0101 0.0053 1.0000
15.750 1.2240 0.08265 0.07997 -0.0135 0.0053 1.0000
16.000 1.2100 0.09016 0.08761 -0.0181 0.0053 1.0000
16.250 1.1987 0.09727 0.09482 -0.0222 0.0054 1.0000
16.500 1.1758 0.10698 0.10467 -0.0279 0.0054 1.0000
16.750 1.1650 0.11424 0.11202 -0.0320 0.0055 1.0000
17.000 1.1361 0.12578 0.12369 -0.0385 0.0055 1.0000
17.250 1.1182 0.13508 0.13308 -0.0438 0.0056 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 474 AIRFOIL (goe474-il)