GOE 797 AIRFOIL (goe797-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 797 AIRFOIL (goe797-il) Reynolds number: 200,000 Max Cl/Cd: 75.95 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe797-il-200000.txt Download as CSV file: xf-goe797-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 797 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.0636 0.09149 0.08782 -0.0926 0.9584 0.0984
-9.500 -0.0487 0.08761 0.08393 -0.0956 0.9555 0.1010
-9.250 -0.3001 0.06431 0.06040 -0.1148 0.9553 0.0765
-9.000 -0.2975 0.05893 0.05491 -0.1185 0.9458 0.0754
-8.750 -0.3230 0.04751 0.04305 -0.1253 0.9367 0.0739
-8.500 -0.3515 0.03790 0.03261 -0.1255 0.9231 0.0736
-8.250 -0.3504 0.03328 0.02735 -0.1243 0.9128 0.0744
-8.000 -0.3338 0.02998 0.02338 -0.1241 0.9068 0.0758
-7.750 -0.3100 0.02827 0.02153 -0.1242 0.9012 0.0774
-7.500 -0.2863 0.02768 0.02094 -0.1238 0.8932 0.0791
-7.250 -0.2584 0.02669 0.01982 -0.1240 0.8887 0.0813
-7.000 -0.2363 0.02540 0.01821 -0.1234 0.8816 0.0836
-6.750 -0.2128 0.02406 0.01656 -0.1228 0.8748 0.0857
-6.500 -0.1836 0.02335 0.01590 -0.1232 0.8705 0.0881
-6.250 -0.1591 0.02293 0.01541 -0.1227 0.8634 0.0910
-6.000 -0.1330 0.02224 0.01446 -0.1223 0.8573 0.0945
-5.750 -0.1046 0.02134 0.01353 -0.1225 0.8532 0.0975
-5.500 -0.0787 0.02106 0.01323 -0.1222 0.8467 0.1009
-5.250 -0.0521 0.02066 0.01265 -0.1219 0.8406 0.1054
-5.000 -0.0235 0.01999 0.01198 -0.1221 0.8363 0.1096
-4.750 0.0040 0.01974 0.01170 -0.1220 0.8314 0.1144
-4.500 0.0299 0.01937 0.01120 -0.1217 0.8251 0.1198
-4.250 0.0584 0.01900 0.01088 -0.1218 0.8204 0.1252
-4.000 0.0887 0.01878 0.01048 -0.1220 0.8168 0.1320
-3.750 0.1131 0.01840 0.01021 -0.1217 0.8105 0.1382
-3.500 0.1410 0.01824 0.00996 -0.1216 0.8055 0.1458
-3.250 0.1701 0.01773 0.00947 -0.1218 0.8015 0.1534
-3.000 0.1982 0.01759 0.00927 -0.1219 0.7971 0.1616
-2.750 0.2239 0.01729 0.00902 -0.1216 0.7914 0.1694
-2.500 0.2524 0.01708 0.00879 -0.1217 0.7866 0.1781
-2.250 0.2824 0.01667 0.00836 -0.1220 0.7825 0.1869
-2.000 0.3078 0.01659 0.00831 -0.1216 0.7757 0.1958
-1.750 0.3354 0.01628 0.00803 -0.1214 0.7698 0.2053
-1.500 0.3659 0.01607 0.00774 -0.1217 0.7651 0.2164
-1.250 0.3910 0.01591 0.00771 -0.1213 0.7585 0.2277
-1.000 0.4184 0.01571 0.00755 -0.1212 0.7525 0.2415
-0.750 0.4486 0.01543 0.00729 -0.1214 0.7474 0.2599
-0.500 0.4735 0.01528 0.00729 -0.1209 0.7405 0.2861
-0.250 0.4998 0.01489 0.00720 -0.1207 0.7346 0.3469
0.000 0.5221 0.01389 0.00720 -0.1197 0.7304 0.6080
0.250 0.5403 0.01341 0.00752 -0.1162 0.7256 0.8808
0.500 0.5958 0.01349 0.00761 -0.1214 0.7199 0.9778
0.750 0.6536 0.01352 0.00750 -0.1276 0.7149 1.0000
1.250 0.6996 0.01375 0.00757 -0.1258 0.7036 1.0000
1.500 0.7238 0.01382 0.00755 -0.1251 0.6977 1.0000
1.750 0.7517 0.01387 0.00746 -0.1250 0.6929 1.0000
2.000 0.7721 0.01405 0.00765 -0.1236 0.6861 1.0000
2.250 0.7969 0.01414 0.00768 -0.1230 0.6799 1.0000
2.500 0.8258 0.01418 0.00760 -0.1230 0.6750 1.0000
2.750 0.8469 0.01436 0.00780 -0.1218 0.6676 1.0000
3.000 0.8730 0.01441 0.00780 -0.1214 0.6610 1.0000
3.250 0.9010 0.01447 0.00777 -0.1213 0.6551 1.0000
3.500 0.9230 0.01458 0.00791 -0.1202 0.6465 1.0000
3.750 0.9518 0.01457 0.00780 -0.1202 0.6398 1.0000
4.000 0.9738 0.01470 0.00796 -0.1191 0.6308 1.0000
4.250 1.0012 0.01470 0.00789 -0.1188 0.6231 1.0000
4.500 1.0239 0.01481 0.00802 -0.1179 0.6136 1.0000
4.750 1.0507 0.01482 0.00796 -0.1175 0.6050 1.0000
5.000 1.0724 0.01494 0.00810 -0.1163 0.5938 1.0000
5.250 1.0973 0.01499 0.00810 -0.1157 0.5833 1.0000
5.500 1.1201 0.01505 0.00811 -0.1146 0.5704 1.0000
5.750 1.1403 0.01517 0.00822 -0.1131 0.5554 1.0000
6.000 1.1605 0.01533 0.00833 -0.1116 0.5397 1.0000
6.250 1.1803 0.01554 0.00848 -0.1101 0.5237 1.0000
6.500 1.1996 0.01581 0.00868 -0.1085 0.5079 1.0000
6.750 1.2185 0.01613 0.00892 -0.1069 0.4929 1.0000
7.000 1.2377 0.01650 0.00921 -0.1054 0.4795 1.0000
7.250 1.2568 0.01691 0.00950 -0.1039 0.4669 1.0000
7.500 1.2736 0.01733 0.00989 -0.1020 0.4540 1.0000
7.750 1.2914 0.01777 0.01030 -0.1004 0.4433 1.0000
8.000 1.3108 0.01824 0.01067 -0.0991 0.4339 1.0000
8.250 1.3260 0.01869 0.01115 -0.0970 0.4248 1.0000
8.500 1.3461 0.01919 0.01155 -0.0958 0.4169 1.0000
8.750 1.3611 0.01967 0.01211 -0.0938 0.4090 1.0000
9.000 1.3802 0.02019 0.01257 -0.0925 0.4017 1.0000
9.250 1.3970 0.02072 0.01314 -0.0909 0.3945 1.0000
9.500 1.4132 0.02127 0.01370 -0.0893 0.3872 1.0000
9.750 1.4329 0.02184 0.01423 -0.0882 0.3804 1.0000
10.000 1.4457 0.02243 0.01491 -0.0861 0.3733 1.0000
10.250 1.4657 0.02302 0.01542 -0.0852 0.3661 1.0000
10.500 1.4762 0.02371 0.01623 -0.0828 0.3590 1.0000
10.750 1.4901 0.02438 0.01691 -0.0810 0.3517 1.0000
11.000 1.5033 0.02512 0.01768 -0.0792 0.3442 1.0000
11.250 1.5125 0.02592 0.01856 -0.0769 0.3364 1.0000
11.500 1.5254 0.02673 0.01936 -0.0752 0.3286 1.0000
11.750 1.5309 0.02773 0.02046 -0.0726 0.3201 1.0000
12.000 1.5410 0.02868 0.02139 -0.0707 0.3118 1.0000
12.250 1.5443 0.02988 0.02271 -0.0681 0.3026 1.0000
12.500 1.5499 0.03110 0.02396 -0.0659 0.2933 1.0000
12.750 1.5526 0.03250 0.02542 -0.0636 0.2834 1.0000
13.000 1.5543 0.03409 0.02710 -0.0614 0.2725 1.0000
13.250 1.5556 0.03578 0.02878 -0.0594 0.2619 1.0000
13.500 1.5554 0.03768 0.03071 -0.0574 0.2505 1.0000
13.750 1.5561 0.03967 0.03277 -0.0557 0.2388 1.0000
14.000 1.5565 0.04178 0.03488 -0.0541 0.2281 1.0000
14.250 1.5554 0.04409 0.03713 -0.0526 0.2184 1.0000
14.500 1.5568 0.04631 0.03941 -0.0514 0.2084 1.0000
14.750 1.5572 0.04864 0.04172 -0.0502 0.2002 1.0000
15.000 1.5574 0.05106 0.04413 -0.0490 0.1923 1.0000
15.250 1.5585 0.05340 0.04649 -0.0480 0.1854 1.0000
15.500 1.5595 0.05581 0.04891 -0.0470 0.1788 1.0000
15.750 1.5612 0.05812 0.05119 -0.0460 0.1729 1.0000
16.000 1.5628 0.06061 0.05378 -0.0454 0.1670 1.0000
16.250 1.5646 0.06290 0.05597 -0.0444 0.1613 1.0000
16.500 1.5650 0.06562 0.05884 -0.0440 0.1561 1.0000
16.750 1.5651 0.06835 0.06163 -0.0435 0.1509 1.0000
17.000 1.5676 0.07066 0.06389 -0.0428 0.1461 1.0000
17.250 1.5665 0.07375 0.06716 -0.0427 0.1416 1.0000
17.500 1.5657 0.07672 0.07020 -0.0426 0.1372 1.0000
17.750 1.5678 0.07915 0.07258 -0.0422 0.1330 1.0000
18.000 1.5643 0.08270 0.07634 -0.0424 0.1289 1.0000
18.250 1.5615 0.08613 0.07987 -0.0427 0.1248 1.0000
18.500 1.5619 0.08890 0.08257 -0.0426 0.1208 1.0000
18.750 1.5563 0.09295 0.08686 -0.0434 0.1169 1.0000
19.000 1.5514 0.09684 0.09086 -0.0441 0.1129 1.0000
19.250 1.5484 0.10026 0.09424 -0.0446 0.1090 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 797 AIRFOIL (goe797-il)