GOE 596 AIRFOIL (goe596-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 596 AIRFOIL (goe596-il) Reynolds number: 500,000 Max Cl/Cd: 105.69 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe596-il-500000.txt Download as CSV file: xf-goe596-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 596 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3437 0.09819 0.09581 -0.0319 1.0000 0.0262
-9.250 -0.3410 0.09531 0.09295 -0.0326 1.0000 0.0268
-9.000 -0.3397 0.09231 0.08997 -0.0334 1.0000 0.0275
-8.750 -0.2945 0.07580 0.07370 -0.0399 1.0000 0.0294
-8.500 -0.2977 0.07348 0.07143 -0.0373 1.0000 0.0297
-8.250 -0.2870 0.07036 0.06831 -0.0382 0.9984 0.0300
-8.000 -0.3824 0.07644 0.07431 -0.0429 0.9969 0.0295
-7.750 -0.3639 0.07314 0.07100 -0.0450 0.9938 0.0298
-7.500 -0.3429 0.06987 0.06772 -0.0488 0.9893 0.0302
-7.250 -0.3185 0.06606 0.06388 -0.0544 0.9850 0.0308
-7.000 -0.2944 0.06167 0.05945 -0.0609 0.9789 0.0317
-6.750 -0.2527 0.04983 0.04713 -0.0795 0.9713 0.0353
-6.500 -0.2379 0.04366 0.04088 -0.0830 0.9636 0.0360
-6.250 -0.2106 0.04136 0.03854 -0.0855 0.9585 0.0366
-6.000 -0.1854 0.03911 0.03622 -0.0874 0.9505 0.0373
-5.750 -0.1581 0.03646 0.03342 -0.0896 0.9429 0.0387
-5.500 -0.1303 0.03393 0.03026 -0.0903 0.9309 0.0427
-5.250 -0.1171 0.02831 0.02443 -0.0907 0.9195 0.0440
-5.000 -0.0933 0.02680 0.02286 -0.0908 0.9096 0.0449
-4.750 -0.0711 0.02548 0.02143 -0.0903 0.8977 0.0460
-4.500 -0.0485 0.02409 0.01986 -0.0897 0.8859 0.0479
-4.250 -0.0200 0.02498 0.02035 -0.0885 0.8746 0.0519
-4.000 -0.0036 0.02056 0.01562 -0.0878 0.8631 0.0540
-3.750 0.0199 0.01949 0.01451 -0.0873 0.8502 0.0555
-3.500 0.0442 0.01870 0.01360 -0.0866 0.8369 0.0581
-3.250 0.0694 0.01523 0.00922 -0.0843 0.8245 0.0434
-3.000 0.0951 0.01433 0.00821 -0.0836 0.8101 0.0421
-2.750 0.1200 0.01326 0.00697 -0.0828 0.7931 0.0417
-2.500 0.1453 0.01273 0.00626 -0.0820 0.7722 0.0421
-2.250 0.1706 0.01238 0.00575 -0.0812 0.7490 0.0425
-2.000 0.1957 0.01189 0.00510 -0.0803 0.7242 0.0427
-1.750 0.2205 0.01132 0.00440 -0.0795 0.7015 0.0429
-1.500 0.2448 0.01068 0.00366 -0.0787 0.6811 0.0436
-1.250 0.2693 0.01030 0.00319 -0.0779 0.6642 0.0444
-1.000 0.2945 0.01007 0.00289 -0.0772 0.6498 0.0454
-0.750 0.3201 0.00991 0.00268 -0.0767 0.6375 0.0469
-0.500 0.3460 0.00981 0.00252 -0.0762 0.6269 0.0490
-0.250 0.3722 0.00970 0.00237 -0.0757 0.6174 0.0508
0.000 0.3985 0.00961 0.00222 -0.0753 0.6090 0.0525
0.250 0.4246 0.00945 0.00204 -0.0748 0.6006 0.0570
0.500 0.4512 0.00940 0.00195 -0.0744 0.5929 0.0632
0.750 0.4733 0.00858 0.00194 -0.0736 0.5856 0.3436
1.000 0.5519 0.00708 0.00210 -0.0847 0.5773 1.0000
1.250 0.5764 0.00717 0.00212 -0.0839 0.5702 1.0000
1.500 0.6009 0.00729 0.00216 -0.0831 0.5636 1.0000
1.750 0.6256 0.00736 0.00220 -0.0824 0.5558 1.0000
2.000 0.6500 0.00749 0.00224 -0.0816 0.5485 1.0000
2.250 0.6749 0.00756 0.00230 -0.0809 0.5407 1.0000
2.500 0.6993 0.00770 0.00236 -0.0801 0.5332 1.0000
2.750 0.7243 0.00777 0.00243 -0.0794 0.5248 1.0000
3.000 0.7488 0.00790 0.00251 -0.0787 0.5170 1.0000
3.250 0.7738 0.00798 0.00259 -0.0780 0.5083 1.0000
3.500 0.7983 0.00811 0.00268 -0.0773 0.4994 1.0000
3.750 0.8226 0.00823 0.00276 -0.0765 0.4880 1.0000
4.000 0.8470 0.00833 0.00285 -0.0757 0.4750 1.0000
4.250 0.8714 0.00845 0.00295 -0.0749 0.4619 1.0000
4.500 0.8955 0.00860 0.00305 -0.0741 0.4467 1.0000
4.750 0.9197 0.00875 0.00318 -0.0734 0.4326 1.0000
5.000 0.9437 0.00893 0.00332 -0.0726 0.4176 1.0000
5.250 0.9671 0.00915 0.00348 -0.0717 0.3992 1.0000
5.500 0.9901 0.00940 0.00366 -0.0708 0.3811 1.0000
5.750 1.0131 0.00968 0.00387 -0.0699 0.3622 1.0000
6.000 1.0359 0.00996 0.00411 -0.0690 0.3442 1.0000
6.250 1.0580 0.01031 0.00437 -0.0680 0.3234 1.0000
6.500 1.0799 0.01068 0.00466 -0.0670 0.3018 1.0000
6.750 1.1008 0.01111 0.00498 -0.0659 0.2737 1.0000
7.000 1.1185 0.01180 0.00541 -0.0643 0.2203 1.0000
7.250 1.1293 0.01303 0.00618 -0.0617 0.1487 1.0000
7.500 1.1342 0.01469 0.00728 -0.0582 0.0637 1.0000
7.750 1.1515 0.01536 0.00792 -0.0565 0.0540 1.0000
8.000 1.1679 0.01607 0.00863 -0.0547 0.0489 1.0000
8.250 1.1870 0.01654 0.00917 -0.0533 0.0466 1.0000
8.500 1.2040 0.01712 0.00978 -0.0516 0.0440 1.0000
8.750 1.2178 0.01786 0.01055 -0.0494 0.0418 1.0000
9.000 1.2272 0.01870 0.01145 -0.0465 0.0399 1.0000
9.250 1.2410 0.01922 0.01205 -0.0443 0.0387 1.0000
9.500 1.2533 0.01984 0.01272 -0.0419 0.0371 1.0000
9.750 1.2639 0.02057 0.01349 -0.0394 0.0358 1.0000
10.000 1.2715 0.02148 0.01445 -0.0367 0.0345 1.0000
10.250 1.2710 0.02292 0.01595 -0.0331 0.0331 1.0000
10.500 1.2792 0.02394 0.01705 -0.0309 0.0322 1.0000
10.750 1.2903 0.02481 0.01800 -0.0291 0.0311 1.0000
11.000 1.2991 0.02589 0.01915 -0.0273 0.0301 1.0000
11.250 1.3073 0.02705 0.02036 -0.0256 0.0290 1.0000
11.500 1.3134 0.02844 0.02180 -0.0238 0.0282 1.0000
11.750 1.3146 0.03030 0.02371 -0.0220 0.0273 1.0000
12.000 1.3122 0.03262 0.02610 -0.0200 0.0265 1.0000
12.250 1.3228 0.03385 0.02743 -0.0191 0.0257 1.0000
12.500 1.3305 0.03540 0.02908 -0.0182 0.0249 1.0000
12.750 1.3367 0.03715 0.03090 -0.0174 0.0241 1.0000
13.000 1.3422 0.03901 0.03283 -0.0166 0.0235 1.0000
13.250 1.3469 0.04101 0.03488 -0.0160 0.0229 1.0000
13.500 1.3501 0.04322 0.03714 -0.0155 0.0223 1.0000
13.750 1.3487 0.04592 0.03988 -0.0143 0.0216 1.0000
14.000 1.3532 0.04813 0.04221 -0.0140 0.0212 1.0000
14.250 1.3577 0.05039 0.04460 -0.0140 0.0206 1.0000
14.500 1.3613 0.05280 0.04711 -0.0139 0.0201 1.0000
14.750 1.3642 0.05534 0.04974 -0.0140 0.0195 1.0000
15.000 1.3668 0.05798 0.05247 -0.0143 0.0190 1.0000
15.250 1.3688 0.06072 0.05530 -0.0146 0.0187 1.0000
15.500 1.3705 0.06357 0.05820 -0.0151 0.0183 1.0000
15.750 1.3716 0.06648 0.06118 -0.0154 0.0179 1.0000
16.000 1.3711 0.06950 0.06424 -0.0151 0.0174 1.0000
16.250 1.3692 0.07299 0.06788 -0.0158 0.0172 1.0000
16.500 1.3669 0.07670 0.07174 -0.0168 0.0169 1.0000
16.750 1.3644 0.08048 0.07567 -0.0179 0.0166 1.0000
17.000 1.3610 0.08448 0.07981 -0.0190 0.0164 1.0000
17.250 1.3566 0.08870 0.08418 -0.0205 0.0161 1.0000
17.500 1.3526 0.09299 0.08859 -0.0222 0.0158 1.0000
17.750 1.3467 0.09766 0.09341 -0.0240 0.0156 1.0000
18.000 1.3418 0.10234 0.09820 -0.0263 0.0153 1.0000
18.250 1.3353 0.10731 0.10329 -0.0286 0.0151 1.0000
18.500 1.3293 0.11230 0.10840 -0.0310 0.0149 1.0000
18.750 1.3234 0.11737 0.11357 -0.0336 0.0147 1.0000
19.000 1.3205 0.12190 0.11815 -0.0361 0.0145 1.0000
19.250 1.3129 0.12738 0.12375 -0.0390 0.0143 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 596 AIRFOIL (goe596-il)