GOE 515 AIRFOIL (goe515-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 515 AIRFOIL (goe515-il) Reynolds number: 500,000 Max Cl/Cd: 103.3 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe515-il-500000.txt Download as CSV file: xf-goe515-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 515 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4738 0.10125 0.09882 -0.0294 1.0000 0.0234
-10.500 -0.7700 0.03916 0.03649 -0.0639 1.0000 0.0172
-10.250 -0.7850 0.03427 0.03123 -0.0623 1.0000 0.0175
-10.000 -0.7897 0.03119 0.02783 -0.0595 1.0000 0.0178
-9.750 -0.7870 0.02914 0.02548 -0.0568 1.0000 0.0182
-9.500 -0.7873 0.02667 0.02269 -0.0535 1.0000 0.0186
-9.250 -0.7861 0.02436 0.02015 -0.0503 1.0000 0.0192
-9.000 -0.7612 0.02397 0.01978 -0.0504 0.9989 0.0201
-8.750 -0.7287 0.02307 0.01874 -0.0522 0.9964 0.0212
-8.500 -0.6955 0.02205 0.01749 -0.0541 0.9941 0.0227
-8.250 -0.6645 0.02102 0.01620 -0.0553 0.9908 0.0237
-8.000 -0.6369 0.01876 0.01373 -0.0566 0.9872 0.0254
-7.750 -0.6024 0.01819 0.01309 -0.0584 0.9849 0.0270
-7.500 -0.5670 0.01764 0.01239 -0.0603 0.9833 0.0290
-7.250 -0.5385 0.01678 0.01133 -0.0607 0.9786 0.0304
-7.000 -0.5087 0.01535 0.00978 -0.0616 0.9752 0.0326
-6.750 -0.4741 0.01487 0.00924 -0.0633 0.9730 0.0349
-6.500 -0.4385 0.01436 0.00862 -0.0650 0.9713 0.0369
-6.250 -0.4028 0.01363 0.00775 -0.0668 0.9700 0.0385
-6.000 -0.3763 0.01273 0.00679 -0.0667 0.9643 0.0412
-5.750 -0.3431 0.01226 0.00629 -0.0679 0.9612 0.0436
-5.500 -0.3088 0.01183 0.00577 -0.0693 0.9585 0.0460
-5.250 -0.2748 0.01128 0.00515 -0.0706 0.9561 0.0484
-5.000 -0.2486 0.01078 0.00465 -0.0702 0.9494 0.0521
-4.750 -0.2177 0.01045 0.00427 -0.0708 0.9447 0.0562
-4.500 -0.1863 0.01005 0.00389 -0.0714 0.9410 0.0634
-4.250 -0.1607 0.00981 0.00369 -0.0708 0.9334 0.0742
-4.000 -0.1311 0.00967 0.00356 -0.0710 0.9280 0.0870
-3.750 -0.1027 0.00969 0.00352 -0.0709 0.9218 0.0952
-3.500 -0.0759 0.00949 0.00335 -0.0706 0.9150 0.1032
-3.250 -0.0474 0.00945 0.00323 -0.0706 0.9093 0.1089
-3.000 -0.0219 0.00922 0.00300 -0.0700 0.9017 0.1145
-2.500 0.0320 0.00899 0.00272 -0.0693 0.8882 0.1237
-2.250 0.0590 0.00880 0.00251 -0.0690 0.8817 0.1295
-2.000 0.0844 0.00868 0.00239 -0.0682 0.8712 0.1362
-1.750 0.1101 0.00854 0.00223 -0.0675 0.8599 0.1427
-1.500 0.1363 0.00842 0.00209 -0.0670 0.8501 0.1502
-1.250 0.1623 0.00830 0.00198 -0.0664 0.8414 0.1584
-1.000 0.1889 0.00819 0.00189 -0.0660 0.8340 0.1706
-0.750 0.2147 0.00803 0.00183 -0.0655 0.8257 0.1950
-0.500 0.2405 0.00785 0.00178 -0.0650 0.8181 0.2396
-0.250 0.2656 0.00762 0.00174 -0.0643 0.8098 0.3015
0.000 0.2880 0.00714 0.00174 -0.0633 0.8016 0.4527
0.250 0.3484 0.00582 0.00191 -0.0702 0.7953 0.9520
0.500 0.3938 0.00595 0.00198 -0.0737 0.7871 0.9800
0.750 0.4437 0.00603 0.00198 -0.0783 0.7757 0.9931
1.000 0.4907 0.00605 0.00193 -0.0824 0.7624 1.0000
1.250 0.5146 0.00607 0.00193 -0.0815 0.7499 1.0000
1.500 0.5384 0.00610 0.00192 -0.0805 0.7366 1.0000
1.750 0.5622 0.00615 0.00191 -0.0794 0.7216 1.0000
2.000 0.5857 0.00621 0.00191 -0.0784 0.7054 1.0000
2.250 0.6092 0.00629 0.00193 -0.0773 0.6882 1.0000
2.500 0.6327 0.00638 0.00197 -0.0763 0.6715 1.0000
2.750 0.6559 0.00649 0.00203 -0.0752 0.6513 1.0000
3.000 0.6787 0.00664 0.00210 -0.0740 0.6299 1.0000
3.250 0.7014 0.00679 0.00219 -0.0728 0.6057 1.0000
3.500 0.7227 0.00702 0.00228 -0.0713 0.5718 1.0000
3.750 0.7433 0.00729 0.00240 -0.0697 0.5305 1.0000
4.000 0.7627 0.00765 0.00257 -0.0680 0.4815 1.0000
4.250 0.7824 0.00804 0.00276 -0.0663 0.4384 1.0000
4.500 0.8025 0.00842 0.00299 -0.0648 0.4020 1.0000
4.750 0.8226 0.00881 0.00323 -0.0633 0.3650 1.0000
5.000 0.8421 0.00925 0.00350 -0.0617 0.3226 1.0000
5.250 0.8606 0.00979 0.00379 -0.0600 0.2693 1.0000
5.500 0.8766 0.01053 0.00416 -0.0579 0.1957 1.0000
5.750 0.8944 0.01116 0.00458 -0.0561 0.1574 1.0000
6.000 0.9145 0.01160 0.00495 -0.0547 0.1365 1.0000
6.250 0.9346 0.01205 0.00531 -0.0533 0.1176 1.0000
6.500 0.9550 0.01247 0.00566 -0.0519 0.0977 1.0000
6.750 0.9735 0.01304 0.00606 -0.0503 0.0687 1.0000
7.000 0.9911 0.01369 0.00661 -0.0485 0.0495 1.0000
7.250 1.0086 0.01435 0.00721 -0.0466 0.0371 1.0000
7.500 1.0260 0.01502 0.00788 -0.0447 0.0306 1.0000
7.750 1.0436 0.01565 0.00851 -0.0429 0.0269 1.0000
8.000 1.0601 0.01636 0.00928 -0.0408 0.0246 1.0000
8.250 1.0779 0.01694 0.00992 -0.0391 0.0228 1.0000
8.500 1.0947 0.01758 0.01057 -0.0373 0.0210 1.0000
8.750 1.1047 0.01873 0.01178 -0.0343 0.0197 1.0000
9.000 1.1209 0.01938 0.01251 -0.0324 0.0190 1.0000
9.250 1.1356 0.02013 0.01334 -0.0302 0.0182 1.0000
9.500 1.1485 0.02090 0.01418 -0.0278 0.0174 1.0000
9.750 1.1599 0.02169 0.01502 -0.0252 0.0167 1.0000
10.000 1.1700 0.02261 0.01599 -0.0225 0.0161 1.0000
10.250 1.1756 0.02416 0.01760 -0.0194 0.0154 1.0000
10.500 1.1856 0.02541 0.01897 -0.0169 0.0149 1.0000
10.750 1.1975 0.02639 0.02006 -0.0148 0.0145 1.0000
11.000 1.2084 0.02755 0.02134 -0.0127 0.0140 1.0000
11.250 1.2185 0.02880 0.02270 -0.0106 0.0136 1.0000
11.500 1.2275 0.03029 0.02433 -0.0085 0.0134 1.0000
11.750 1.2359 0.03156 0.02571 -0.0065 0.0129 1.0000
12.000 1.2427 0.03321 0.02748 -0.0045 0.0127 1.0000
12.250 1.2486 0.03477 0.02915 -0.0026 0.0125 1.0000
12.500 1.2531 0.03652 0.03101 -0.0008 0.0122 1.0000
12.750 1.2553 0.03871 0.03331 0.0009 0.0119 1.0000
13.000 1.2540 0.04147 0.03623 0.0027 0.0118 1.0000
13.250 1.2508 0.04432 0.03927 0.0043 0.0117 1.0000
13.500 1.2420 0.04793 0.04310 0.0058 0.0116 1.0000
13.750 1.2270 0.05232 0.04774 0.0068 0.0115 1.0000
14.000 1.2158 0.05612 0.05174 0.0072 0.0115 1.0000
14.250 1.1960 0.06135 0.05721 0.0069 0.0115 1.0000
14.500 1.1783 0.06648 0.06254 0.0058 0.0115 1.0000
14.750 1.1648 0.07132 0.06755 0.0042 0.0115 1.0000
15.000 1.1505 0.07662 0.07301 0.0020 0.0115 1.0000
15.250 1.1269 0.08402 0.08061 -0.0017 0.0115 1.0000
15.500 1.1076 0.09136 0.08813 -0.0059 0.0115 1.0000
15.750 1.0875 0.09941 0.09633 -0.0108 0.0115 1.0000
16.000 1.0804 0.10534 0.10237 -0.0146 0.0116 1.0000
16.250 1.0614 0.11429 0.11148 -0.0204 0.0117 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 515 AIRFOIL (goe515-il)