RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 25 AIRFOIL (raf25-il) Reynolds number: 1,000,000 Max Cl/Cd: 72.23 at α=1.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf25-il-1000000.txt Download as CSV file: xf-raf25-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 25 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4869 0.07482 0.07328 -0.0155 1.0000 0.0122
-8.750 -0.4920 0.06972 0.06819 -0.0173 1.0000 0.0122
-8.500 -0.4992 0.06429 0.06279 -0.0194 1.0000 0.0123
-8.250 -0.5083 0.05884 0.05735 -0.0221 1.0000 0.0123
-8.000 -0.5416 0.05235 0.05090 -0.0268 1.0000 0.0113
-7.750 -0.5548 0.04538 0.04385 -0.0316 1.0000 0.0107
-7.500 -0.6418 0.03850 0.03628 -0.0340 1.0000 0.0071
-7.250 -0.6681 0.02227 0.01882 -0.0294 1.0000 0.0071
-7.000 -0.6531 0.01954 0.01572 -0.0276 1.0000 0.0075
-6.750 -0.6344 0.01785 0.01376 -0.0261 1.0000 0.0077
-6.500 -0.6142 0.01659 0.01230 -0.0248 1.0000 0.0079
-6.250 -0.5963 0.01448 0.00984 -0.0231 1.0000 0.0081
-6.000 -0.5793 0.01222 0.00726 -0.0212 1.0000 0.0089
-5.750 -0.5566 0.01181 0.00683 -0.0202 1.0000 0.0097
-5.500 -0.5338 0.01143 0.00640 -0.0193 1.0000 0.0104
-5.250 -0.5116 0.01089 0.00578 -0.0181 1.0000 0.0113
-5.000 -0.4892 0.01048 0.00531 -0.0170 1.0000 0.0120
-4.750 -0.4674 0.00997 0.00472 -0.0158 1.0000 0.0126
-4.500 -0.4367 0.00910 0.00375 -0.0166 0.9987 0.0150
-4.250 -0.4005 0.00900 0.00367 -0.0185 0.9967 0.0175
-4.000 -0.3645 0.00874 0.00338 -0.0204 0.9947 0.0195
-3.750 -0.3317 0.00824 0.00286 -0.0216 0.9921 0.0243
-3.500 -0.2979 0.00809 0.00269 -0.0229 0.9892 0.0273
-3.250 -0.2637 0.00781 0.00238 -0.0244 0.9864 0.0302
-3.000 -0.2290 0.00747 0.00201 -0.0260 0.9841 0.0348
-2.750 -0.1964 0.00726 0.00178 -0.0270 0.9804 0.0388
-2.500 -0.1643 0.00705 0.00155 -0.0280 0.9754 0.0425
-2.250 -0.1301 0.00677 0.00135 -0.0294 0.9718 0.0588
-2.000 -0.0991 0.00617 0.00114 -0.0303 0.9657 0.1660
-1.750 -0.0669 0.00560 0.00099 -0.0316 0.9584 0.2902
-1.500 -0.0337 0.00515 0.00087 -0.0330 0.9472 0.3923
-1.250 0.0004 0.00476 0.00078 -0.0345 0.9305 0.4979
-1.000 0.0290 0.00448 0.00072 -0.0347 0.9106 0.5860
-0.750 0.0553 0.00433 0.00069 -0.0342 0.8938 0.6452
-0.500 0.0801 0.00418 0.00068 -0.0334 0.8764 0.7068
-0.250 0.1023 0.00395 0.00068 -0.0320 0.8577 0.7922
0.000 0.1251 0.00368 0.00071 -0.0305 0.8348 0.9074
0.250 0.1877 0.00379 0.00078 -0.0380 0.7981 0.9739
0.500 0.2219 0.00391 0.00080 -0.0394 0.7736 0.9856
0.750 0.2587 0.00405 0.00083 -0.0413 0.7411 0.9936
1.000 0.2976 0.00426 0.00083 -0.0438 0.6859 0.9991
1.250 0.3236 0.00448 0.00084 -0.0434 0.6274 1.0000
1.500 0.3450 0.00479 0.00087 -0.0421 0.5553 1.0000
1.750 0.3666 0.00516 0.00094 -0.0408 0.4759 1.0000
2.000 0.3867 0.00572 0.00106 -0.0394 0.3594 1.0000
2.250 0.4058 0.00649 0.00128 -0.0379 0.2125 1.0000
2.500 0.4243 0.00740 0.00158 -0.0363 0.0603 1.0000
2.750 0.4484 0.00765 0.00178 -0.0355 0.0494 1.0000
3.000 0.4732 0.00781 0.00195 -0.0349 0.0451 1.0000
3.250 0.4978 0.00799 0.00215 -0.0342 0.0407 1.0000
3.500 0.5216 0.00833 0.00250 -0.0333 0.0341 1.0000
3.750 0.5471 0.00841 0.00258 -0.0328 0.0311 1.0000
4.000 0.5711 0.00871 0.00284 -0.0320 0.0232 1.0000
4.250 0.5959 0.00891 0.00301 -0.0313 0.0190 1.0000
4.500 0.6183 0.00949 0.00368 -0.0301 0.0155 1.0000
4.750 0.6428 0.00975 0.00396 -0.0294 0.0142 1.0000
5.000 0.6669 0.01005 0.00428 -0.0286 0.0127 1.0000
5.250 0.6898 0.01056 0.00481 -0.0276 0.0111 1.0000
5.500 0.7080 0.01183 0.00626 -0.0256 0.0100 1.0000
5.750 0.7318 0.01225 0.00672 -0.0248 0.0096 1.0000
6.000 0.7547 0.01285 0.00742 -0.0238 0.0091 1.0000
6.250 0.7772 0.01357 0.00822 -0.0228 0.0084 1.0000
6.500 0.7993 0.01445 0.00920 -0.0216 0.0079 1.0000
6.750 0.8213 0.01542 0.01028 -0.0205 0.0075 1.0000
7.000 0.8436 0.01621 0.01114 -0.0197 0.0071 1.0000
7.250 0.8620 0.01817 0.01332 -0.0181 0.0066 1.0000
7.500 0.8755 0.02169 0.01730 -0.0158 0.0064 1.0000
7.750 0.8890 0.02481 0.02080 -0.0136 0.0064 1.0000
8.000 0.8942 0.02955 0.02606 -0.0107 0.0064 1.0000
8.250 0.9016 0.03330 0.03017 -0.0083 0.0064 1.0000
8.500 0.8877 0.04126 0.03869 -0.0043 0.0064 1.0000
8.750 0.9022 0.04306 0.04061 -0.0031 0.0065 1.0000
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