GOE 391 AIRFOIL (goe391-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 391 AIRFOIL (goe391-il) Reynolds number: 1,000,000 Max Cl/Cd: 104.7 at α=1.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe391-il-1000000.txt Download as CSV file: xf-goe391-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 391 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4436 0.08840 0.08684 -0.0180 1.0000 0.0074
-7.500 -0.4549 0.08646 0.08494 -0.0161 1.0000 0.0074
-7.250 -0.4589 0.08378 0.08228 -0.0161 1.0000 0.0074
-7.000 -0.4472 0.07966 0.07815 -0.0204 0.9991 0.0074
-6.750 -0.4223 0.07398 0.07244 -0.0285 0.9969 0.0075
-6.500 -0.4054 0.06778 0.06621 -0.0344 0.9951 0.0080
-6.250 -0.3811 0.06399 0.06238 -0.0394 0.9933 0.0083
-6.000 -0.3557 0.06017 0.05852 -0.0444 0.9907 0.0092
-5.750 -0.3240 0.05551 0.05378 -0.0504 0.9882 0.0106
-5.500 -0.2842 0.05099 0.04913 -0.0568 0.9862 0.0113
-5.250 -0.2495 0.04652 0.04452 -0.0616 0.9846 0.0114
-5.000 -0.2148 0.04187 0.03973 -0.0661 0.9833 0.0114
-4.750 -0.1881 0.03788 0.03559 -0.0679 0.9793 0.0114
-4.500 -0.1587 0.03376 0.03130 -0.0699 0.9762 0.0114
-4.000 -0.0988 0.02203 0.01885 -0.0739 0.9720 0.0104
-3.750 -0.0656 0.01687 0.01319 -0.0745 0.9705 0.0113
-3.500 -0.0454 0.01118 0.00667 -0.0726 0.9652 0.0124
-3.250 -0.0138 0.01154 0.00707 -0.0735 0.9618 0.0135
-3.000 0.0173 0.01080 0.00621 -0.0741 0.9591 0.0154
-2.750 0.0484 0.01014 0.00541 -0.0747 0.9567 0.0166
-2.500 0.0728 0.00890 0.00402 -0.0739 0.9517 0.0184
-2.250 0.1013 0.00898 0.00414 -0.0741 0.9471 0.0204
-2.000 0.1310 0.00865 0.00376 -0.0744 0.9430 0.0226
-1.750 0.1567 0.00839 0.00346 -0.0738 0.9359 0.0243
-1.500 0.1857 0.00855 0.00359 -0.0739 0.9300 0.0253
-1.250 0.2087 0.00734 0.00233 -0.0729 0.9223 0.0292
-1.000 0.2359 0.00705 0.00201 -0.0727 0.9150 0.0307
-0.750 0.2611 0.00678 0.00171 -0.0720 0.9051 0.0316
-0.500 0.2873 0.00656 0.00146 -0.0715 0.8962 0.0331
-0.250 0.3140 0.00639 0.00124 -0.0712 0.8877 0.0349
0.000 0.3400 0.00628 0.00109 -0.0706 0.8746 0.0363
0.250 0.3652 0.00622 0.00097 -0.0699 0.8528 0.0369
0.500 0.3905 0.00614 0.00079 -0.0692 0.8342 0.0383
0.750 0.4157 0.00612 0.00069 -0.0685 0.8116 0.0395
1.000 0.4394 0.00620 0.00062 -0.0675 0.7731 0.0416
1.250 0.4603 0.00642 0.00059 -0.0658 0.7047 0.0471
1.500 0.5256 0.00502 0.00095 -0.0755 0.5716 1.0000
1.750 0.5454 0.00540 0.00103 -0.0738 0.5027 1.0000
2.000 0.5622 0.00605 0.00117 -0.0717 0.3790 1.0000
2.250 0.5822 0.00651 0.00133 -0.0702 0.3112 1.0000
2.500 0.6038 0.00686 0.00146 -0.0690 0.2593 1.0000
2.750 0.6219 0.00754 0.00171 -0.0673 0.1576 1.0000
3.000 0.6415 0.00813 0.00196 -0.0657 0.0792 1.0000
3.250 0.6632 0.00854 0.00220 -0.0645 0.0420 1.0000
3.500 0.6869 0.00875 0.00241 -0.0637 0.0382 1.0000
3.750 0.7102 0.00901 0.00267 -0.0627 0.0351 1.0000
4.000 0.7335 0.00929 0.00300 -0.0617 0.0326 1.0000
4.250 0.7565 0.00961 0.00337 -0.0607 0.0311 1.0000
4.500 0.7784 0.01006 0.00389 -0.0595 0.0295 1.0000
4.750 0.8014 0.01036 0.00423 -0.0586 0.0288 1.0000
5.000 0.8257 0.01051 0.00438 -0.0579 0.0280 1.0000
5.250 0.8494 0.01074 0.00465 -0.0572 0.0267 1.0000
5.500 0.8729 0.01101 0.00494 -0.0564 0.0247 1.0000
5.750 0.8966 0.01126 0.00519 -0.0556 0.0221 1.0000
6.000 0.9161 0.01200 0.00600 -0.0541 0.0187 1.0000
6.250 0.9419 0.01200 0.00600 -0.0538 0.0173 1.0000
6.500 0.9671 0.01209 0.00608 -0.0534 0.0150 1.0000
6.750 0.9858 0.01298 0.00703 -0.0517 0.0122 1.0000
7.000 1.0094 0.01325 0.00734 -0.0509 0.0115 1.0000
7.250 1.0332 0.01349 0.00758 -0.0503 0.0103 1.0000
7.500 1.0562 0.01384 0.00793 -0.0495 0.0091 1.0000
7.750 1.0725 0.01509 0.00932 -0.0474 0.0080 1.0000
8.000 1.0882 0.01654 0.01096 -0.0453 0.0075 1.0000
8.250 1.1088 0.01728 0.01183 -0.0441 0.0073 1.0000
8.500 1.1295 0.01798 0.01263 -0.0429 0.0068 1.0000
8.750 1.1481 0.01905 0.01385 -0.0415 0.0065 1.0000
9.000 1.1667 0.02013 0.01507 -0.0400 0.0061 1.0000
9.250 1.1863 0.02084 0.01587 -0.0388 0.0056 1.0000
9.500 1.2036 0.02194 0.01711 -0.0373 0.0053 1.0000
9.750 1.2159 0.02384 0.01922 -0.0351 0.0049 1.0000
10.000 1.2113 0.02860 0.02456 -0.0306 0.0046 1.0000
10.250 1.1975 0.03400 0.03054 -0.0253 0.0045 1.0000
10.500 1.1799 0.03884 0.03581 -0.0200 0.0045 1.0000
10.750 1.1142 0.02975 0.02703 -0.0108 0.0046 1.0000
11.000 1.0978 0.03262 0.03008 -0.0072 0.0046 1.0000
11.250 1.0730 0.03669 0.03435 -0.0043 0.0046 1.0000
11.500 1.0505 0.04119 0.03902 -0.0031 0.0046 1.0000
11.750 1.0323 0.04591 0.04387 -0.0033 0.0046 1.0000
12.000 1.0108 0.05214 0.05024 -0.0051 0.0046 1.0000
12.250 0.9953 0.05843 0.05664 -0.0080 0.0046 1.0000
12.500 0.9769 0.06647 0.06478 -0.0124 0.0047 1.0000
12.750 0.9592 0.07566 0.07408 -0.0177 0.0047 1.0000
13.000 0.9467 0.08440 0.08290 -0.0227 0.0048 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 391 AIRFOIL (goe391-il)