GOE 595 AIRFOIL (goe595-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 595 AIRFOIL (goe595-il) Reynolds number: 500,000 Max Cl/Cd: 97.73 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe595-il-500000.txt Download as CSV file: xf-goe595-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 595 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4283 0.08357 0.08145 -0.0338 1.0000 0.0194
-8.250 -0.4447 0.08129 0.07923 -0.0318 1.0000 0.0195
-8.000 -0.4660 0.07935 0.07735 -0.0293 0.9999 0.0196
-7.750 -0.4503 0.07121 0.06919 -0.0425 0.9948 0.0205
-7.500 -0.4280 0.06139 0.05922 -0.0559 0.9889 0.0215
-7.250 -0.4075 0.05293 0.05046 -0.0641 0.9849 0.0218
-7.000 -0.3651 0.02659 0.02373 -0.0720 0.9738 0.0224
-6.750 -0.3477 0.02136 0.01832 -0.0744 0.9717 0.0231
-6.500 -0.3290 0.01928 0.01618 -0.0748 0.9669 0.0238
-6.250 -0.3078 0.01622 0.01287 -0.0758 0.9628 0.0246
-6.000 -0.3246 0.02567 0.02143 -0.0727 0.9623 0.0204
-5.750 -0.2956 0.02228 0.01757 -0.0737 0.9599 0.0204
-5.500 -0.2636 0.01924 0.01407 -0.0749 0.9583 0.0201
-5.250 -0.2345 0.01744 0.01196 -0.0751 0.9551 0.0205
-5.000 -0.2063 0.01635 0.01065 -0.0750 0.9507 0.0213
-4.750 -0.1725 0.01514 0.00920 -0.0762 0.9485 0.0223
-4.500 -0.1403 0.01347 0.00751 -0.0774 0.9463 0.0248
-4.250 -0.1051 0.01259 0.00658 -0.0790 0.9444 0.0266
-4.000 -0.0787 0.01196 0.00588 -0.0785 0.9387 0.0282
-3.750 -0.0479 0.01151 0.00536 -0.0790 0.9341 0.0298
-3.500 -0.0175 0.01041 0.00419 -0.0795 0.9302 0.0334
-3.250 0.0070 0.00998 0.00371 -0.0787 0.9232 0.0360
-3.000 0.0364 0.00960 0.00327 -0.0789 0.9182 0.0390
-2.750 0.0652 0.00927 0.00288 -0.0790 0.9133 0.0436
-2.500 0.0908 0.00901 0.00264 -0.0784 0.9065 0.0511
-2.250 0.1193 0.00862 0.00242 -0.0785 0.9014 0.0993
-2.000 0.1419 0.00816 0.00229 -0.0775 0.8938 0.1945
-1.750 0.1683 0.00774 0.00214 -0.0773 0.8872 0.2922
-1.500 0.1882 0.00713 0.00205 -0.0759 0.8776 0.4529
-1.250 0.2099 0.00664 0.00198 -0.0745 0.8691 0.5904
-1.000 0.2295 0.00620 0.00193 -0.0725 0.8594 0.7153
-0.750 0.2501 0.00586 0.00195 -0.0705 0.8494 0.8277
-0.500 0.2904 0.00574 0.00199 -0.0727 0.8425 0.9169
-0.250 0.3449 0.00581 0.00203 -0.0782 0.8327 0.9589
0.000 0.3972 0.00592 0.00208 -0.0832 0.8216 0.9776
0.250 0.4400 0.00597 0.00205 -0.0863 0.8081 0.9861
0.500 0.4870 0.00603 0.00203 -0.0905 0.7924 0.9944
0.750 0.5318 0.00604 0.00194 -0.0942 0.7738 0.9998
1.000 0.5533 0.00609 0.00189 -0.0929 0.7528 1.0000
1.250 0.5739 0.00616 0.00186 -0.0913 0.7308 1.0000
1.500 0.5944 0.00626 0.00185 -0.0897 0.7077 1.0000
1.750 0.6146 0.00638 0.00186 -0.0881 0.6828 1.0000
2.000 0.6344 0.00652 0.00189 -0.0864 0.6554 1.0000
2.250 0.6538 0.00669 0.00195 -0.0846 0.6251 1.0000
2.500 0.6727 0.00690 0.00202 -0.0827 0.5924 1.0000
2.750 0.6914 0.00714 0.00212 -0.0808 0.5604 1.0000
3.000 0.7089 0.00746 0.00224 -0.0787 0.5226 1.0000
3.250 0.7274 0.00776 0.00240 -0.0768 0.4910 1.0000
3.500 0.7474 0.00801 0.00255 -0.0752 0.4698 1.0000
3.750 0.7686 0.00822 0.00270 -0.0739 0.4524 1.0000
4.000 0.7894 0.00845 0.00287 -0.0725 0.4326 1.0000
4.250 0.8106 0.00868 0.00305 -0.0712 0.4153 1.0000
4.500 0.8316 0.00891 0.00323 -0.0699 0.3958 1.0000
4.750 0.8526 0.00915 0.00341 -0.0685 0.3713 1.0000
5.000 0.8729 0.00944 0.00360 -0.0671 0.3411 1.0000
5.250 0.8907 0.00988 0.00384 -0.0652 0.2935 1.0000
5.500 0.9075 0.01043 0.00417 -0.0633 0.2516 1.0000
5.750 0.9255 0.01093 0.00452 -0.0615 0.2204 1.0000
6.000 0.9431 0.01150 0.00489 -0.0597 0.1796 1.0000
6.250 0.9476 0.01308 0.00574 -0.0558 0.0598 1.0000
6.500 0.9655 0.01368 0.00633 -0.0541 0.0481 1.0000
6.750 0.9845 0.01419 0.00690 -0.0524 0.0426 1.0000
7.000 1.0035 0.01469 0.00744 -0.0509 0.0380 1.0000
7.250 1.0185 0.01548 0.00828 -0.0486 0.0329 1.0000
7.500 1.0378 0.01593 0.00879 -0.0471 0.0301 1.0000
7.750 1.0548 0.01652 0.00942 -0.0453 0.0269 1.0000
8.000 1.0640 0.01762 0.01060 -0.0421 0.0242 1.0000
8.250 1.0835 0.01797 0.01100 -0.0407 0.0223 1.0000
8.500 1.0974 0.01857 0.01166 -0.0384 0.0207 1.0000
8.750 1.1102 0.01922 0.01233 -0.0359 0.0194 1.0000
9.000 1.1112 0.02064 0.01383 -0.0315 0.0180 1.0000
9.250 1.1210 0.02160 0.01488 -0.0287 0.0174 1.0000
9.500 1.1345 0.02231 0.01569 -0.0266 0.0166 1.0000
9.750 1.1470 0.02310 0.01655 -0.0244 0.0156 1.0000
10.000 1.1588 0.02395 0.01746 -0.0223 0.0148 1.0000
10.250 1.1705 0.02476 0.01831 -0.0203 0.0140 1.0000
10.500 1.1771 0.02616 0.01978 -0.0177 0.0134 1.0000
10.750 1.1802 0.02882 0.02257 -0.0148 0.0128 1.0000
11.000 1.1910 0.02999 0.02388 -0.0129 0.0125 1.0000
11.250 1.2010 0.03142 0.02546 -0.0110 0.0122 1.0000
11.500 1.2101 0.03319 0.02738 -0.0091 0.0119 1.0000
11.750 1.2179 0.03506 0.02942 -0.0072 0.0116 1.0000
12.000 1.2235 0.03697 0.03150 -0.0053 0.0113 1.0000
12.250 1.2267 0.03942 0.03416 -0.0033 0.0111 1.0000
12.500 1.2291 0.04101 0.03585 -0.0016 0.0106 1.0000
12.750 1.2294 0.04328 0.03827 0.0001 0.0104 1.0000
13.000 1.2310 0.04459 0.03963 0.0012 0.0101 1.0000
13.250 1.2273 0.04747 0.04270 0.0025 0.0100 1.0000
13.500 1.2210 0.05074 0.04616 0.0036 0.0099 1.0000
13.750 1.2168 0.05345 0.04894 0.0039 0.0096 1.0000
14.000 1.2062 0.05745 0.05312 0.0042 0.0095 1.0000
14.250 1.1932 0.06194 0.05782 0.0039 0.0095 1.0000
14.500 1.1776 0.06701 0.06308 0.0029 0.0094 1.0000
14.750 1.1626 0.07226 0.06853 0.0013 0.0095 1.0000
15.000 1.1437 0.07857 0.07503 -0.0012 0.0094 1.0000
15.250 1.1274 0.08486 0.08150 -0.0042 0.0095 1.0000
15.500 1.1032 0.09324 0.09014 -0.0085 0.0097 1.0000
15.750 1.0726 0.10352 0.10067 -0.0143 0.0098 1.0000
16.000 0.9664 0.13496 0.13263 -0.0333 0.0114 1.0000
16.250 0.9232 0.15400 0.15175 -0.0439 0.0120 1.0000
16.500 0.9073 0.16352 0.16124 -0.0485 0.0128 1.0000
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Polar data table (+)
Polar graphs
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