GOE 564 AIRFOIL (goe564-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 564 AIRFOIL (goe564-il) Reynolds number: 500,000 Max Cl/Cd: 95.67 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe564-il-500000.txt Download as CSV file: xf-goe564-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 564 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3711 0.08674 0.08453 -0.0349 1.0000 0.0284
-9.500 -0.3756 0.08239 0.08020 -0.0355 1.0000 0.0284
-8.250 -0.5620 0.06436 0.06224 -0.0378 1.0000 0.0228
-8.000 -0.5656 0.06194 0.05982 -0.0371 1.0000 0.0230
-7.750 -0.6169 0.02708 0.02320 -0.0523 0.9932 0.0210
-7.500 -0.5893 0.02494 0.02060 -0.0535 0.9898 0.0219
-7.250 -0.5649 0.02174 0.01706 -0.0548 0.9870 0.0235
-7.000 -0.5319 0.02127 0.01654 -0.0564 0.9852 0.0248
-6.750 -0.4996 0.02000 0.01504 -0.0579 0.9838 0.0262
-6.500 -0.4727 0.01897 0.01378 -0.0579 0.9797 0.0275
-6.250 -0.4433 0.01766 0.01217 -0.0586 0.9767 0.0289
-6.000 -0.4137 0.01621 0.01062 -0.0595 0.9746 0.0310
-5.750 -0.3805 0.01552 0.00983 -0.0609 0.9729 0.0326
-5.500 -0.3461 0.01480 0.00900 -0.0623 0.9716 0.0343
-5.250 -0.3173 0.01435 0.00845 -0.0625 0.9679 0.0358
-5.000 -0.2884 0.01352 0.00748 -0.0628 0.9642 0.0371
-4.750 -0.2572 0.01251 0.00640 -0.0636 0.9618 0.0392
-4.500 -0.2231 0.01203 0.00590 -0.0650 0.9601 0.0415
-4.250 -0.1883 0.01157 0.00539 -0.0664 0.9587 0.0434
-4.000 -0.1529 0.01114 0.00491 -0.0680 0.9577 0.0452
-3.750 -0.1289 0.01079 0.00452 -0.0671 0.9522 0.0469
-3.500 -0.0974 0.01030 0.00404 -0.0679 0.9493 0.0512
-3.250 -0.0626 0.00997 0.00371 -0.0693 0.9474 0.0557
-3.000 -0.0273 0.00960 0.00340 -0.0708 0.9458 0.0691
-2.750 0.0081 0.00932 0.00324 -0.0724 0.9442 0.0988
-2.500 0.0374 0.00920 0.00314 -0.0726 0.9400 0.1121
-2.250 0.0670 0.00903 0.00300 -0.0728 0.9353 0.1220
-2.000 0.1006 0.00880 0.00279 -0.0739 0.9320 0.1313
-1.750 0.1341 0.00858 0.00257 -0.0749 0.9269 0.1405
-1.500 0.1632 0.00831 0.00234 -0.0749 0.9178 0.1515
-1.250 0.1922 0.00807 0.00214 -0.0749 0.9085 0.1688
-1.000 0.2224 0.00776 0.00196 -0.0752 0.9007 0.2039
-0.750 0.2445 0.00710 0.00187 -0.0742 0.8901 0.3764
-0.500 0.2607 0.00607 0.00185 -0.0718 0.8792 0.6859
-0.250 0.2921 0.00549 0.00192 -0.0717 0.8683 0.9136
0.000 0.3537 0.00551 0.00188 -0.0785 0.8525 0.9666
0.250 0.3981 0.00558 0.00187 -0.0818 0.8369 0.9837
0.500 0.4436 0.00563 0.00184 -0.0855 0.8230 0.9932
0.750 0.4885 0.00567 0.00180 -0.0892 0.8074 1.0000
1.000 0.5108 0.00571 0.00177 -0.0879 0.7890 1.0000
1.250 0.5330 0.00577 0.00175 -0.0865 0.7681 1.0000
1.500 0.5541 0.00586 0.00173 -0.0849 0.7391 1.0000
1.750 0.5740 0.00600 0.00171 -0.0830 0.6999 1.0000
2.000 0.5923 0.00625 0.00174 -0.0808 0.6561 1.0000
2.250 0.6115 0.00652 0.00182 -0.0788 0.6224 1.0000
2.500 0.6323 0.00673 0.00192 -0.0772 0.5949 1.0000
2.750 0.6533 0.00695 0.00202 -0.0757 0.5659 1.0000
3.000 0.6737 0.00719 0.00214 -0.0741 0.5291 1.0000
3.250 0.6919 0.00754 0.00226 -0.0720 0.4737 1.0000
3.500 0.7082 0.00806 0.00246 -0.0697 0.4039 1.0000
3.750 0.7251 0.00863 0.00271 -0.0676 0.3360 1.0000
4.000 0.7432 0.00916 0.00297 -0.0657 0.2807 1.0000
4.250 0.7637 0.00955 0.00321 -0.0643 0.2514 1.0000
4.500 0.7855 0.00985 0.00344 -0.0631 0.2326 1.0000
4.750 0.8076 0.01014 0.00366 -0.0619 0.2175 1.0000
5.000 0.8299 0.01041 0.00390 -0.0608 0.2044 1.0000
5.250 0.8524 0.01069 0.00414 -0.0598 0.1922 1.0000
5.500 0.8749 0.01096 0.00437 -0.0587 0.1796 1.0000
5.750 0.8976 0.01123 0.00462 -0.0577 0.1659 1.0000
6.000 0.9203 0.01150 0.00486 -0.0567 0.1501 1.0000
6.250 0.9423 0.01184 0.00511 -0.0557 0.1259 1.0000
6.500 0.9625 0.01234 0.00544 -0.0543 0.0983 1.0000
6.750 0.9827 0.01284 0.00587 -0.0530 0.0817 1.0000
7.000 1.0019 0.01346 0.00637 -0.0514 0.0537 1.0000
7.250 1.0198 0.01422 0.00698 -0.0497 0.0370 1.0000
7.500 1.0376 0.01501 0.00778 -0.0478 0.0311 1.0000
7.750 1.0577 0.01555 0.00840 -0.0464 0.0285 1.0000
8.000 1.0757 0.01627 0.00917 -0.0447 0.0258 1.0000
8.250 1.0902 0.01729 0.01026 -0.0424 0.0238 1.0000
8.500 1.1092 0.01787 0.01092 -0.0409 0.0226 1.0000
8.750 1.1265 0.01858 0.01170 -0.0391 0.0214 1.0000
9.000 1.1432 0.01931 0.01248 -0.0373 0.0202 1.0000
9.250 1.1554 0.02042 0.01362 -0.0348 0.0189 1.0000
9.500 1.1664 0.02169 0.01500 -0.0322 0.0180 1.0000
9.750 1.1822 0.02233 0.01573 -0.0302 0.0173 1.0000
10.000 1.1950 0.02322 0.01671 -0.0278 0.0167 1.0000
10.250 1.2075 0.02410 0.01768 -0.0255 0.0160 1.0000
10.500 1.2190 0.02510 0.01876 -0.0231 0.0155 1.0000
10.750 1.2302 0.02605 0.01977 -0.0208 0.0150 1.0000
11.000 1.2393 0.02735 0.02115 -0.0183 0.0145 1.0000
11.250 1.2453 0.02948 0.02338 -0.0158 0.0140 1.0000
11.500 1.2520 0.03129 0.02536 -0.0133 0.0136 1.0000
11.750 1.2597 0.03237 0.02659 -0.0110 0.0133 1.0000
12.000 1.2652 0.03394 0.02833 -0.0085 0.0130 1.0000
12.250 1.2694 0.03555 0.03010 -0.0062 0.0126 1.0000
12.500 1.2718 0.03731 0.03201 -0.0040 0.0123 1.0000
12.750 1.2721 0.03942 0.03429 -0.0019 0.0121 1.0000
13.000 1.2717 0.04148 0.03649 -0.0001 0.0118 1.0000
13.250 1.2691 0.04391 0.03908 0.0015 0.0117 1.0000
13.500 1.2645 0.04663 0.04198 0.0029 0.0116 1.0000
13.750 1.2579 0.04965 0.04516 0.0037 0.0114 1.0000
14.000 1.2515 0.05277 0.04842 0.0041 0.0113 1.0000
14.250 1.2444 0.05615 0.05194 0.0039 0.0111 1.0000
14.500 1.2314 0.06074 0.05673 0.0031 0.0111 1.0000
14.750 1.2139 0.06634 0.06253 0.0014 0.0111 1.0000
15.000 1.1966 0.07233 0.06872 -0.0011 0.0111 1.0000
15.250 1.1826 0.07814 0.07466 -0.0041 0.0110 1.0000
15.500 1.1609 0.08598 0.08271 -0.0085 0.0111 1.0000
15.750 1.1381 0.09473 0.09166 -0.0139 0.0111 1.0000
16.000 1.1067 0.10621 0.10335 -0.0212 0.0113 1.0000
16.250 1.0177 0.13449 0.13206 -0.0392 0.0123 1.0000
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