GOE 484 AIRFOIL (goe484-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 484 AIRFOIL (goe484-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.73 at α=1.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe484-il-1000000-n5.txt Download as CSV file: xf-goe484-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 484 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3343 0.10186 0.10009 -0.0456 0.9944 0.0092
-9.750 -0.3240 0.09843 0.09665 -0.0482 0.9926 0.0097
-9.500 -0.3217 0.09212 0.09035 -0.0527 0.9901 0.0109
-9.250 -0.3064 0.08932 0.08755 -0.0558 0.9881 0.0111
-9.000 -0.2904 0.08632 0.08454 -0.0592 0.9863 0.0113
-8.750 -0.2776 0.08372 0.08194 -0.0616 0.9825 0.0115
-8.500 -0.2641 0.08052 0.07874 -0.0649 0.9790 0.0118
-8.250 -0.2487 0.07718 0.07539 -0.0689 0.9761 0.0124
-7.750 -0.2214 0.06625 0.06443 -0.0829 0.9627 0.0142
-7.500 -0.1938 0.06317 0.06133 -0.0892 0.9588 0.0144
-7.250 -0.1656 0.05984 0.05797 -0.0958 0.9546 0.0148
-7.000 -0.1380 0.05608 0.05418 -0.1028 0.9471 0.0153
-6.750 -0.1080 0.04301 0.04091 -0.1227 0.9309 0.0175
-6.500 -0.0822 0.03792 0.03568 -0.1297 0.9199 0.0179
-6.250 -0.0570 0.03505 0.03271 -0.1331 0.9104 0.0182
-6.000 -0.0313 0.03236 0.02989 -0.1359 0.9014 0.0185
-5.750 -0.0107 0.01765 0.01420 -0.1443 0.8882 0.0198
-5.500 0.0138 0.01552 0.01169 -0.1447 0.8809 0.0204
-5.250 0.0393 0.01415 0.01004 -0.1448 0.8745 0.0209
-5.000 0.0652 0.01319 0.00887 -0.1448 0.8674 0.0213
-4.500 0.1179 0.01197 0.00731 -0.1445 0.8539 0.0217
-4.250 0.1444 0.01146 0.00666 -0.1443 0.8472 0.0218
-4.000 0.1711 0.01105 0.00614 -0.1441 0.8414 0.0219
-3.750 0.1979 0.01066 0.00564 -0.1439 0.8355 0.0220
-3.500 0.2244 0.01008 0.00494 -0.1437 0.8302 0.0222
-3.250 0.2509 0.00938 0.00413 -0.1435 0.8252 0.0223
-3.000 0.2775 0.00885 0.00352 -0.1433 0.8194 0.0225
-2.500 0.3311 0.00819 0.00278 -0.1429 0.8084 0.0232
-2.250 0.3581 0.00794 0.00250 -0.1426 0.8026 0.0234
-2.000 0.3850 0.00773 0.00224 -0.1423 0.7971 0.0236
-1.750 0.4121 0.00753 0.00203 -0.1421 0.7907 0.0239
-1.500 0.4389 0.00738 0.00185 -0.1418 0.7838 0.0243
-1.250 0.4660 0.00723 0.00168 -0.1416 0.7760 0.0247
-1.000 0.4926 0.00713 0.00154 -0.1412 0.7667 0.0252
-0.750 0.5190 0.00705 0.00142 -0.1408 0.7537 0.0257
-0.500 0.5446 0.00701 0.00132 -0.1402 0.7321 0.0262
-0.250 0.5658 0.00721 0.00125 -0.1386 0.6708 0.0266
0.250 0.6118 0.00759 0.00127 -0.1365 0.5930 0.0272
0.500 0.6375 0.00765 0.00126 -0.1360 0.5762 0.0275
0.750 0.6633 0.00771 0.00126 -0.1355 0.5603 0.0277
1.000 0.6887 0.00781 0.00129 -0.1349 0.5409 0.0280
1.250 0.7142 0.00788 0.00129 -0.1344 0.5226 0.0290
1.500 0.7391 0.00802 0.00132 -0.1337 0.4947 0.0300
1.750 0.7632 0.00823 0.00138 -0.1329 0.4575 0.0314
2.000 0.7816 0.00893 0.00162 -0.1312 0.3501 0.0333
2.500 0.8239 0.00991 0.00215 -0.1288 0.2165 0.1051
3.000 0.8635 0.01125 0.00292 -0.1260 0.0469 0.1746
3.250 0.8881 0.01109 0.00317 -0.1256 0.0424 0.3942
3.750 0.9398 0.01041 0.00370 -0.1251 0.0367 1.0000
4.000 0.9650 0.01061 0.00389 -0.1245 0.0358 1.0000
4.250 0.9899 0.01082 0.00409 -0.1239 0.0349 1.0000
4.500 1.0147 0.01104 0.00431 -0.1233 0.0338 1.0000
4.750 1.0392 0.01128 0.00454 -0.1227 0.0326 1.0000
5.000 1.0635 0.01154 0.00479 -0.1220 0.0316 1.0000
5.250 1.0873 0.01183 0.00508 -0.1212 0.0304 1.0000
5.500 1.1105 0.01218 0.00545 -0.1203 0.0289 1.0000
5.750 1.1336 0.01253 0.00583 -0.1194 0.0280 1.0000
6.000 1.1576 0.01276 0.00606 -0.1187 0.0277 1.0000
6.250 1.1814 0.01301 0.00632 -0.1179 0.0272 1.0000
6.500 1.2049 0.01327 0.00661 -0.1171 0.0265 1.0000
6.750 1.2282 0.01354 0.00689 -0.1163 0.0256 1.0000
7.000 1.2512 0.01382 0.00719 -0.1155 0.0247 1.0000
7.250 1.2741 0.01410 0.00746 -0.1146 0.0236 1.0000
7.500 1.2964 0.01442 0.00779 -0.1137 0.0224 1.0000
7.750 1.3170 0.01487 0.00825 -0.1124 0.0211 1.0000
8.000 1.3396 0.01514 0.00854 -0.1115 0.0206 1.0000
8.250 1.3617 0.01542 0.00884 -0.1106 0.0196 1.0000
8.500 1.3834 0.01571 0.00915 -0.1095 0.0182 1.0000
8.750 1.4045 0.01604 0.00947 -0.1084 0.0167 1.0000
9.000 1.4245 0.01644 0.00988 -0.1071 0.0154 1.0000
9.250 1.4437 0.01685 0.01031 -0.1057 0.0143 1.0000
9.500 1.4615 0.01727 0.01073 -0.1040 0.0129 1.0000
9.750 1.4779 0.01775 0.01122 -0.1020 0.0116 1.0000
10.000 1.4940 0.01823 0.01176 -0.1001 0.0110 1.0000
10.250 1.5094 0.01875 0.01232 -0.0980 0.0103 1.0000
10.500 1.5243 0.01932 0.01293 -0.0959 0.0096 1.0000
10.750 1.5382 0.01996 0.01360 -0.0937 0.0090 1.0000
11.000 1.5515 0.02065 0.01434 -0.0915 0.0084 1.0000
11.250 1.5653 0.02129 0.01507 -0.0894 0.0082 1.0000
11.500 1.5786 0.02199 0.01583 -0.0873 0.0078 1.0000
11.750 1.5910 0.02275 0.01665 -0.0851 0.0075 1.0000
12.000 1.6029 0.02356 0.01753 -0.0830 0.0072 1.0000
12.250 1.6139 0.02445 0.01848 -0.0808 0.0068 1.0000
12.500 1.6241 0.02543 0.01951 -0.0786 0.0065 1.0000
12.750 1.6319 0.02662 0.02079 -0.0762 0.0061 1.0000
13.000 1.6426 0.02760 0.02185 -0.0743 0.0060 1.0000
13.250 1.6520 0.02870 0.02304 -0.0723 0.0058 1.0000
13.500 1.6604 0.02991 0.02434 -0.0703 0.0057 1.0000
13.750 1.6684 0.03119 0.02571 -0.0684 0.0055 1.0000
14.000 1.6752 0.03260 0.02721 -0.0665 0.0054 1.0000
14.250 1.6817 0.03408 0.02879 -0.0647 0.0052 1.0000
14.500 1.6874 0.03567 0.03047 -0.0630 0.0051 1.0000
14.750 1.6919 0.03743 0.03234 -0.0613 0.0050 1.0000
15.000 1.6955 0.03933 0.03433 -0.0598 0.0049 1.0000
15.250 1.6984 0.04136 0.03647 -0.0583 0.0047 1.0000
15.500 1.6994 0.04366 0.03887 -0.0570 0.0046 1.0000
15.750 1.6989 0.04621 0.04153 -0.0559 0.0045 1.0000
16.000 1.6965 0.04910 0.04454 -0.0549 0.0044 1.0000
16.250 1.6914 0.05245 0.04803 -0.0542 0.0042 1.0000
16.500 1.6878 0.05578 0.05148 -0.0539 0.0042 1.0000
16.750 1.6848 0.05919 0.05503 -0.0539 0.0042 1.0000
17.000 1.6793 0.06308 0.05906 -0.0542 0.0041 1.0000
17.250 1.6745 0.06706 0.06317 -0.0548 0.0041 1.0000
17.500 1.6671 0.07162 0.06787 -0.0558 0.0041 1.0000
17.750 1.6577 0.07666 0.07306 -0.0572 0.0040 1.0000
18.000 1.6479 0.08194 0.07848 -0.0589 0.0040 1.0000
18.250 1.6352 0.08794 0.08465 -0.0612 0.0040 1.0000
18.500 1.6205 0.09455 0.09141 -0.0640 0.0040 1.0000
18.750 1.6041 0.10174 0.09875 -0.0674 0.0039 1.0000
19.000 1.5846 0.10991 0.10709 -0.0716 0.0039 1.0000
19.250 1.5627 0.11905 0.11640 -0.0767 0.0039 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 484 AIRFOIL (goe484-il)