GOE 491 AIRFOIL (goe491-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 491 AIRFOIL (goe491-il) Reynolds number: 50,000 Max Cl/Cd: 37.1 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe491-il-50000-n5.txt Download as CSV file: xf-goe491-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 491 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4755 0.11483 0.10782 -0.0067 1.0000 0.0680
-8.500 -0.4774 0.11276 0.10585 -0.0090 1.0000 0.0683
-8.250 -0.4780 0.11045 0.10362 -0.0113 1.0000 0.0685
-8.000 -0.4753 0.10781 0.10103 -0.0144 1.0000 0.0686
-7.750 -0.4614 0.10062 0.09389 -0.0103 1.0000 0.0700
-7.500 -0.4540 0.09661 0.08991 -0.0100 1.0000 0.0714
-7.250 -0.4481 0.09323 0.08658 -0.0109 1.0000 0.0725
-7.000 -0.4423 0.08990 0.08325 -0.0121 1.0000 0.0737
-6.750 -0.4357 0.08660 0.07998 -0.0136 1.0000 0.0750
-6.500 -0.4283 0.08340 0.07681 -0.0155 1.0000 0.0777
-6.250 -0.4186 0.08079 0.07417 -0.0198 1.0000 0.0813
-6.000 -0.4051 0.07826 0.07151 -0.0243 1.0000 0.0823
-5.750 -0.3990 0.07317 0.06658 -0.0215 1.0000 0.0851
-5.500 -0.3876 0.06971 0.06312 -0.0217 1.0000 0.0876
-5.250 -0.3606 0.06447 0.05749 -0.0269 1.0000 0.0551
-5.000 -0.3510 0.06054 0.05362 -0.0257 1.0000 0.0523
-4.750 -0.3353 0.05678 0.04976 -0.0262 1.0000 0.0496
-4.500 -0.3170 0.05289 0.04570 -0.0271 1.0000 0.0473
-4.250 -0.2967 0.04893 0.04149 -0.0278 1.0000 0.0455
-4.000 -0.2764 0.04524 0.03751 -0.0279 1.0000 0.0449
-3.750 -0.2584 0.04267 0.03479 -0.0273 1.0000 0.0477
-3.500 -0.2376 0.03978 0.03157 -0.0268 1.0000 0.0507
-3.250 -0.2158 0.03674 0.02813 -0.0262 1.0000 0.0521
-3.000 -0.1932 0.03369 0.02461 -0.0252 1.0000 0.0532
-2.750 -0.1690 0.03075 0.02101 -0.0241 1.0000 0.0558
-2.500 -0.1491 0.02927 0.01943 -0.0231 1.0000 0.0646
-2.250 -0.1245 0.02687 0.01643 -0.0219 1.0000 0.0701
-2.000 -0.1007 0.02536 0.01447 -0.0208 1.0000 0.0828
-1.750 -0.0766 0.02398 0.01276 -0.0200 1.0000 0.0963
-1.500 -0.0524 0.02282 0.01142 -0.0194 1.0000 0.1116
-1.250 -0.0256 0.02180 0.01018 -0.0189 1.0000 0.1203
-1.000 0.0039 0.02097 0.00905 -0.0188 1.0000 0.1265
-0.750 0.0292 0.02044 0.00837 -0.0183 1.0000 0.1376
-0.500 0.0532 0.01992 0.00781 -0.0177 1.0000 0.1456
-0.250 0.0760 0.01956 0.00732 -0.0168 1.0000 0.1532
0.000 0.0977 0.01924 0.00703 -0.0159 1.0000 0.1640
0.250 0.1240 0.01896 0.00682 -0.0161 0.9968 0.1869
0.500 0.1681 0.01812 0.00655 -0.0201 0.9815 0.2832
0.750 0.2358 0.01656 0.00622 -0.0280 0.9711 1.0000
1.000 0.2811 0.01667 0.00614 -0.0317 0.9496 1.0000
1.250 0.3284 0.01672 0.00608 -0.0355 0.9281 1.0000
1.500 0.3749 0.01672 0.00604 -0.0390 0.9043 1.0000
1.750 0.4192 0.01672 0.00602 -0.0419 0.8770 1.0000
2.000 0.4602 0.01672 0.00604 -0.0440 0.8448 1.0000
2.250 0.4956 0.01676 0.00607 -0.0448 0.8068 1.0000
2.500 0.5291 0.01683 0.00611 -0.0451 0.7667 1.0000
2.750 0.5597 0.01698 0.00622 -0.0448 0.7262 1.0000
3.000 0.5871 0.01724 0.00636 -0.0439 0.6870 1.0000
3.250 0.6119 0.01758 0.00660 -0.0426 0.6489 1.0000
3.500 0.6358 0.01797 0.00692 -0.0413 0.6143 1.0000
3.750 0.6587 0.01840 0.00730 -0.0399 0.5818 1.0000
4.000 0.6818 0.01885 0.00773 -0.0386 0.5534 1.0000
4.250 0.7054 0.01934 0.00822 -0.0375 0.5282 1.0000
4.500 0.7292 0.01984 0.00886 -0.0366 0.5035 1.0000
4.750 0.7519 0.02036 0.00942 -0.0353 0.4758 1.0000
5.000 0.7733 0.02091 0.00996 -0.0338 0.4444 1.0000
5.250 0.7941 0.02146 0.01054 -0.0323 0.4121 1.0000
5.500 0.8141 0.02198 0.01125 -0.0307 0.3775 1.0000
5.750 0.8330 0.02245 0.01184 -0.0290 0.3382 1.0000
6.000 0.8508 0.02294 0.01239 -0.0272 0.2875 1.0000
6.250 0.8686 0.02370 0.01310 -0.0255 0.2304 1.0000
6.500 0.8845 0.02513 0.01427 -0.0238 0.1639 1.0000
6.750 0.8993 0.02711 0.01593 -0.0221 0.1158 1.0000
7.000 0.9151 0.02890 0.01777 -0.0205 0.0922 1.0000
7.250 0.9307 0.03082 0.01980 -0.0186 0.0807 1.0000
7.500 0.9457 0.03273 0.02184 -0.0168 0.0679 1.0000
7.750 0.9606 0.03495 0.02418 -0.0150 0.0593 1.0000
8.000 0.9754 0.03709 0.02647 -0.0133 0.0514 1.0000
8.250 0.9922 0.03996 0.02961 -0.0116 0.0469 1.0000
8.500 1.0069 0.04261 0.03259 -0.0100 0.0420 1.0000
8.750 1.0199 0.04599 0.03605 -0.0087 0.0389 1.0000
9.000 1.0296 0.04961 0.04023 -0.0068 0.0376 1.0000
9.250 1.0333 0.05333 0.04451 -0.0048 0.0361 1.0000
9.500 1.0326 0.05708 0.04873 -0.0030 0.0350 1.0000
9.750 1.0273 0.06087 0.05289 -0.0012 0.0340 1.0000
10.000 1.0174 0.06468 0.05701 0.0003 0.0333 1.0000
10.250 1.0019 0.06852 0.06108 0.0018 0.0333 1.0000
10.500 0.9837 0.07293 0.06569 0.0017 0.0336 1.0000
10.750 0.9641 0.07822 0.07114 -0.0004 0.0341 1.0000
11.000 0.9444 0.08470 0.07775 -0.0044 0.0350 1.0000
11.250 0.9261 0.09214 0.08526 -0.0096 0.0357 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 491 AIRFOIL (goe491-il)