GOE 444 AIRFOIL (goe444-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 444 AIRFOIL (goe444-il) Reynolds number: 200,000 Max Cl/Cd: 38.17 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe444-il-200000.txt Download as CSV file: xf-goe444-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 444 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6333 0.08734 0.08386 -0.0107 1.0000 0.0278
-8.250 -0.6388 0.08326 0.07975 -0.0129 1.0000 0.0279
-8.000 -0.6413 0.07922 0.07563 -0.0147 1.0000 0.0280
-7.750 -0.6418 0.07539 0.07169 -0.0158 1.0000 0.0281
-7.500 -0.6406 0.07166 0.06780 -0.0163 1.0000 0.0282
-7.250 -0.6376 0.06798 0.06394 -0.0163 1.0000 0.0283
-7.000 -0.6431 0.06001 0.05590 -0.0169 1.0000 0.0291
-6.750 -0.6361 0.05553 0.05146 -0.0165 1.0000 0.0302
-6.500 -0.6274 0.05211 0.04798 -0.0159 1.0000 0.0319
-6.250 -0.6169 0.04868 0.04439 -0.0152 1.0000 0.0339
-6.000 -0.6043 0.04528 0.04073 -0.0143 1.0000 0.0365
-5.500 -0.5785 0.03815 0.03285 -0.0112 1.0000 0.0431
-5.250 -0.5627 0.03531 0.02991 -0.0101 1.0000 0.0466
-5.000 -0.5471 0.03292 0.02698 -0.0079 1.0000 0.0555
-4.750 -0.5293 0.03018 0.02420 -0.0068 1.0000 0.0618
-4.000 -0.4582 0.02202 0.01453 0.0001 1.0000 0.0391
-3.750 -0.4333 0.01882 0.01093 0.0020 1.0000 0.0315
-3.500 -0.4076 0.01781 0.00969 0.0035 1.0000 0.0288
-3.000 -0.3618 0.01452 0.00625 0.0063 1.0000 0.0289
-2.750 -0.3426 0.01301 0.00476 0.0081 1.0000 0.0325
-2.500 -0.3208 0.01239 0.00411 0.0094 1.0000 0.0415
-2.250 -0.2999 0.01148 0.00312 0.0110 1.0000 0.0556
-2.000 -0.3048 0.00800 0.00256 0.0171 1.0000 0.6977
-1.750 -0.1653 0.00843 0.00323 -0.0030 1.0000 0.9848
-1.500 -0.0955 0.00840 0.00299 -0.0120 1.0000 0.9993
-1.250 -0.0777 0.00822 0.00275 -0.0106 1.0000 1.0000
-1.000 -0.0631 0.00806 0.00255 -0.0084 1.0000 1.0000
-0.750 -0.0480 0.00795 0.00238 -0.0063 1.0000 1.0000
-0.500 -0.0323 0.00788 0.00228 -0.0041 1.0000 1.0000
-0.250 -0.0163 0.00783 0.00222 -0.0021 1.0000 1.0000
0.000 0.0000 0.00782 0.00220 0.0000 1.0000 1.0000
0.250 0.0163 0.00783 0.00222 0.0021 1.0000 1.0000
0.500 0.0324 0.00787 0.00228 0.0042 1.0000 1.0000
0.750 0.0480 0.00795 0.00238 0.0063 1.0000 1.0000
1.000 0.0631 0.00806 0.00255 0.0084 1.0000 1.0000
1.250 0.0777 0.00821 0.00274 0.0106 1.0000 1.0000
1.500 0.0957 0.00840 0.00299 0.0120 0.9993 1.0000
1.750 0.1649 0.00843 0.00323 0.0030 0.9849 1.0000
2.000 0.3050 0.00799 0.00256 -0.0171 0.7020 1.0000
2.250 0.2997 0.01151 0.00314 -0.0110 0.0546 1.0000
2.500 0.3205 0.01244 0.00416 -0.0093 0.0412 1.0000
2.750 0.3425 0.01302 0.00477 -0.0081 0.0330 1.0000
3.000 0.3617 0.01453 0.00626 -0.0063 0.0289 1.0000
3.250 0.3835 0.01644 0.00820 -0.0047 0.0282 1.0000
3.500 0.4076 0.01784 0.00971 -0.0035 0.0288 1.0000
3.750 0.4333 0.01880 0.01091 -0.0020 0.0314 1.0000
4.000 0.4582 0.02200 0.01451 -0.0001 0.0391 1.0000
4.750 0.5292 0.03023 0.02424 0.0068 0.0624 1.0000
5.000 0.5471 0.03296 0.02701 0.0079 0.0554 1.0000
5.250 0.5628 0.03531 0.02992 0.0101 0.0465 1.0000
5.500 0.5785 0.03814 0.03283 0.0112 0.0431 1.0000
6.000 0.6044 0.04526 0.04071 0.0143 0.0365 1.0000
6.250 0.6171 0.04868 0.04439 0.0152 0.0339 1.0000
6.500 0.6275 0.05211 0.04799 0.0159 0.0319 1.0000
6.750 0.6363 0.05554 0.05148 0.0165 0.0303 1.0000
7.000 0.6433 0.05999 0.05588 0.0169 0.0291 1.0000
7.250 0.6379 0.06799 0.06394 0.0163 0.0283 1.0000
7.500 0.6409 0.07168 0.06782 0.0163 0.0282 1.0000
7.750 0.6421 0.07542 0.07171 0.0158 0.0281 1.0000
8.000 0.6414 0.07929 0.07570 0.0146 0.0280 1.0000
8.250 0.6387 0.08332 0.07981 0.0127 0.0279 1.0000
8.500 0.6334 0.08747 0.08399 0.0103 0.0278 1.0000
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Polar data table (+)
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