GOE 360 AIRFOIL (goe360-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 360 AIRFOIL (goe360-il) Reynolds number: 1,000,000 Max Cl/Cd: 131.22 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe360-il-1000000.txt Download as CSV file: xf-goe360-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 360 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1666 0.09083 0.08901 -0.0370 0.8688 0.0150
-9.000 -0.1649 0.08736 0.08551 -0.0385 0.8623 0.0151
-8.750 -0.1647 0.08366 0.08181 -0.0404 0.8549 0.0151
-8.500 -0.1610 0.07954 0.07766 -0.0407 0.8474 0.0153
-8.250 -0.1535 0.07646 0.07456 -0.0407 0.8382 0.0154
-8.000 -0.1472 0.07362 0.07170 -0.0412 0.8288 0.0156
-7.750 -0.2283 0.08368 0.08174 -0.0390 0.8503 0.0153
-7.500 -0.2184 0.08099 0.07904 -0.0403 0.8410 0.0155
-7.250 -0.2057 0.07822 0.07623 -0.0428 0.8311 0.0156
-7.000 -0.1912 0.07532 0.07329 -0.0461 0.8197 0.0159
-6.750 -0.1749 0.07226 0.07018 -0.0498 0.8074 0.0163
-6.500 -0.1568 0.06901 0.06687 -0.0538 0.7936 0.0170
-6.250 -0.1262 0.06410 0.06184 -0.0635 0.7778 0.0183
-6.000 -0.0971 0.05915 0.05674 -0.0711 0.7601 0.0184
-5.750 -0.0728 0.05526 0.05269 -0.0750 0.7398 0.0185
-5.500 -0.0598 0.05177 0.04911 -0.0759 0.7175 0.0188
-5.000 -0.0157 0.04713 0.04422 -0.0795 0.6788 0.0193
-4.750 0.0095 0.04470 0.04166 -0.0817 0.6644 0.0198
-4.500 0.0472 0.04144 0.03817 -0.0857 0.6531 0.0221
-4.250 0.0787 0.03820 0.03472 -0.0880 0.6440 0.0222
-4.000 0.1066 0.03540 0.03174 -0.0894 0.6356 0.0222
-3.750 0.1301 0.03120 0.02737 -0.0911 0.6286 0.0226
-3.500 0.1542 0.02971 0.02579 -0.0918 0.6206 0.0229
-3.250 0.1799 0.02825 0.02425 -0.0924 0.6135 0.0233
-3.000 0.2069 0.02671 0.02258 -0.0931 0.6065 0.0239
-2.750 0.2350 0.02509 0.02081 -0.0936 0.6003 0.0250
-2.500 0.2688 0.02375 0.01919 -0.0933 0.5940 0.0267
-2.000 0.3240 0.01888 0.01384 -0.0940 0.5812 0.0273
-1.750 0.3508 0.01785 0.01271 -0.0943 0.5738 0.0277
-1.500 0.3783 0.01701 0.01179 -0.0945 0.5671 0.0282
-1.250 0.4060 0.01622 0.01089 -0.0947 0.5594 0.0290
-1.000 0.4345 0.01546 0.00998 -0.0947 0.5517 0.0307
-0.750 0.4643 0.01565 0.00996 -0.0943 0.5423 0.0324
-0.500 0.4927 0.01497 0.00911 -0.0943 0.5331 0.0325
0.000 0.5477 0.01240 0.00630 -0.0947 0.5128 0.0346
0.250 0.5756 0.01203 0.00585 -0.0948 0.5013 0.0359
0.500 0.6036 0.01177 0.00546 -0.0947 0.4886 0.0379
0.750 0.6318 0.01229 0.00584 -0.0945 0.4749 0.0396
1.000 0.6590 0.01088 0.00433 -0.0946 0.4624 0.0421
1.250 0.6865 0.01067 0.00406 -0.0946 0.4503 0.0442
1.500 0.7144 0.01022 0.00351 -0.0944 0.4396 0.0422
1.750 0.7416 0.01025 0.00354 -0.0946 0.4290 0.0522
2.250 0.7969 0.00939 0.00249 -0.0941 0.4098 0.0342
2.500 0.8243 0.00936 0.00242 -0.0941 0.4011 0.0350
2.750 0.8520 0.00935 0.00239 -0.0941 0.3931 0.0361
3.000 0.8795 0.00938 0.00239 -0.0941 0.3857 0.0373
3.250 0.9071 0.00941 0.00240 -0.0942 0.3787 0.0385
3.500 0.9344 0.00950 0.00246 -0.0942 0.3717 0.0395
3.750 0.9621 0.00951 0.00247 -0.0942 0.3654 0.0435
4.000 0.9892 0.00962 0.00255 -0.0942 0.3588 0.0481
4.250 1.0128 0.00807 0.00285 -0.0939 0.3541 1.0000
4.500 1.0401 0.00822 0.00295 -0.0939 0.3489 1.0000
4.750 1.0669 0.00841 0.00309 -0.0938 0.3433 1.0000
5.000 1.0943 0.00853 0.00322 -0.0939 0.3392 1.0000
5.250 1.1214 0.00868 0.00335 -0.0939 0.3345 1.0000
5.750 1.1749 0.00903 0.00365 -0.0938 0.3219 1.0000
6.000 1.2013 0.00923 0.00382 -0.0938 0.3151 1.0000
6.250 1.2279 0.00940 0.00399 -0.0937 0.3094 1.0000
6.500 1.2545 0.00957 0.00415 -0.0937 0.3033 1.0000
6.750 1.2804 0.00980 0.00435 -0.0936 0.2966 1.0000
7.000 1.3069 0.00996 0.00453 -0.0936 0.2894 1.0000
7.250 1.3325 0.01021 0.00475 -0.0934 0.2818 1.0000
7.500 1.3584 0.01041 0.00494 -0.0933 0.2729 1.0000
7.750 1.3837 0.01067 0.00518 -0.0931 0.2622 1.0000
8.000 1.4077 0.01104 0.00547 -0.0928 0.2443 1.0000
8.250 1.4273 0.01180 0.00598 -0.0919 0.2037 1.0000
8.500 1.4235 0.01469 0.00799 -0.0878 0.0757 1.0000
8.750 1.4301 0.01646 0.00943 -0.0850 0.0216 1.0000
9.000 1.4497 0.01704 0.01003 -0.0840 0.0188 1.0000
9.250 1.4671 0.01774 0.01078 -0.0827 0.0165 1.0000
9.500 1.4858 0.01829 0.01137 -0.0816 0.0156 1.0000
9.750 1.5024 0.01890 0.01203 -0.0802 0.0147 1.0000
10.000 1.5155 0.01959 0.01276 -0.0782 0.0140 1.0000
10.250 1.5257 0.02049 0.01373 -0.0759 0.0132 1.0000
10.500 1.5312 0.02175 0.01509 -0.0733 0.0125 1.0000
10.750 1.5432 0.02264 0.01604 -0.0718 0.0123 1.0000
11.000 1.5538 0.02369 0.01715 -0.0702 0.0119 1.0000
11.250 1.5633 0.02486 0.01840 -0.0687 0.0115 1.0000
11.500 1.5718 0.02619 0.01978 -0.0673 0.0111 1.0000
11.750 1.5789 0.02767 0.02133 -0.0659 0.0107 1.0000
12.000 1.5838 0.02942 0.02316 -0.0646 0.0104 1.0000
12.250 1.5843 0.03168 0.02551 -0.0632 0.0101 1.0000
12.500 1.5780 0.03477 0.02872 -0.0619 0.0098 1.0000
12.750 1.5644 0.03890 0.03300 -0.0609 0.0096 1.0000
13.000 1.5670 0.04145 0.03563 -0.0606 0.0095 1.0000
13.250 1.5675 0.04430 0.03858 -0.0604 0.0093 1.0000
13.500 1.5658 0.04749 0.04187 -0.0603 0.0092 1.0000
13.750 1.5620 0.05098 0.04546 -0.0603 0.0091 1.0000
14.000 1.5559 0.05481 0.04939 -0.0604 0.0089 1.0000
14.250 1.5494 0.05879 0.05347 -0.0606 0.0088 1.0000
14.500 1.5419 0.06299 0.05777 -0.0610 0.0087 1.0000
14.750 1.5339 0.06731 0.06218 -0.0614 0.0086 1.0000
15.000 1.5255 0.07176 0.06673 -0.0620 0.0084 1.0000
15.250 1.5176 0.07622 0.07128 -0.0626 0.0083 1.0000
15.500 1.5091 0.08082 0.07597 -0.0634 0.0082 1.0000
15.750 1.5018 0.08530 0.08054 -0.0642 0.0081 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 360 AIRFOIL (goe360-il)