RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 25 AIRFOIL (raf25-il) Reynolds number: 500,000 Max Cl/Cd: 68.18 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf25-il-500000.txt Download as CSV file: xf-raf25-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5679 0.08017 0.07793 -0.0179 1.0000 0.0187
-8.250 -0.5730 0.07504 0.07284 -0.0222 1.0000 0.0187
-8.000 -0.5749 0.06867 0.06642 -0.0279 1.0000 0.0187
-7.750 -0.5755 0.06297 0.06062 -0.0310 1.0000 0.0187
-7.500 -0.5751 0.05784 0.05535 -0.0326 1.0000 0.0188
-7.250 -0.5738 0.05279 0.05012 -0.0333 1.0000 0.0188
-7.000 -0.5698 0.04816 0.04525 -0.0332 1.0000 0.0188
-6.250 -0.5625 0.02919 0.02531 -0.0301 1.0000 0.0157
-6.000 -0.5478 0.02563 0.02135 -0.0282 1.0000 0.0161
-5.750 -0.5323 0.02196 0.01720 -0.0261 1.0000 0.0160
-5.500 -0.5149 0.01858 0.01333 -0.0240 1.0000 0.0158
-5.250 -0.4943 0.01666 0.01112 -0.0224 1.0000 0.0166
-5.000 -0.4726 0.01538 0.00963 -0.0211 1.0000 0.0176
-4.750 -0.4500 0.01468 0.00878 -0.0199 1.0000 0.0186
-4.500 -0.4305 0.01271 0.00664 -0.0183 1.0000 0.0208
-4.250 -0.4084 0.01229 0.00620 -0.0173 1.0000 0.0233
-4.000 -0.3866 0.01168 0.00553 -0.0160 1.0000 0.0257
-3.750 -0.3647 0.01133 0.00512 -0.0148 1.0000 0.0277
-3.500 -0.3450 0.01034 0.00404 -0.0133 1.0000 0.0325
-3.250 -0.3232 0.01001 0.00369 -0.0122 0.9999 0.0368
-3.000 -0.2856 0.00972 0.00334 -0.0144 0.9969 0.0401
-2.750 -0.2497 0.00916 0.00276 -0.0163 0.9938 0.0480
-2.500 -0.2144 0.00885 0.00241 -0.0180 0.9898 0.0552
-2.250 -0.1789 0.00833 0.00210 -0.0198 0.9860 0.1024
-2.000 -0.1452 0.00734 0.00192 -0.0218 0.9830 0.3157
-1.750 -0.1153 0.00671 0.00181 -0.0226 0.9766 0.4620
-1.500 -0.0816 0.00622 0.00174 -0.0241 0.9721 0.5837
-1.250 -0.0497 0.00581 0.00168 -0.0249 0.9669 0.6835
-1.000 -0.0206 0.00532 0.00164 -0.0249 0.9604 0.7999
-0.750 0.0247 0.00501 0.00169 -0.0282 0.9590 0.9262
-0.500 0.0973 0.00507 0.00176 -0.0378 0.9630 0.9969
-0.250 0.1423 0.00501 0.00167 -0.0416 0.9566 1.0000
0.000 0.1811 0.00494 0.00157 -0.0440 0.9449 1.0000
0.250 0.2252 0.00484 0.00141 -0.0473 0.9201 1.0000
0.500 0.2579 0.00486 0.00128 -0.0480 0.8805 1.0000
0.750 0.2825 0.00493 0.00125 -0.0470 0.8515 1.0000
1.000 0.3060 0.00502 0.00126 -0.0459 0.8244 1.0000
1.250 0.3291 0.00512 0.00126 -0.0446 0.7949 1.0000
1.500 0.3513 0.00527 0.00127 -0.0432 0.7554 1.0000
1.750 0.3728 0.00547 0.00128 -0.0416 0.7044 1.0000
2.000 0.3934 0.00577 0.00131 -0.0399 0.6345 1.0000
2.250 0.4133 0.00619 0.00141 -0.0382 0.5479 1.0000
2.500 0.4319 0.00685 0.00157 -0.0364 0.4237 1.0000
2.750 0.4459 0.00823 0.00193 -0.0343 0.1797 1.0000
3.000 0.4652 0.00916 0.00232 -0.0328 0.0651 1.0000
3.250 0.4890 0.00950 0.00267 -0.0320 0.0549 1.0000
3.500 0.5120 0.00999 0.00319 -0.0309 0.0463 1.0000
3.750 0.5362 0.01027 0.00351 -0.0302 0.0400 1.0000
4.000 0.5581 0.01095 0.00422 -0.0289 0.0328 1.0000
4.250 0.5826 0.01123 0.00450 -0.0282 0.0272 1.0000
4.500 0.6029 0.01220 0.00552 -0.0267 0.0231 1.0000
4.750 0.6263 0.01272 0.00613 -0.0257 0.0211 1.0000
5.000 0.6490 0.01342 0.00689 -0.0246 0.0192 1.0000
5.250 0.6721 0.01408 0.00757 -0.0236 0.0174 1.0000
5.500 0.6912 0.01614 0.00977 -0.0219 0.0154 1.0000
5.750 0.7148 0.01712 0.01088 -0.0209 0.0148 1.0000
6.000 0.7377 0.01865 0.01262 -0.0198 0.0141 1.0000
6.250 0.7594 0.02087 0.01513 -0.0183 0.0135 1.0000
6.500 0.7778 0.02450 0.01921 -0.0162 0.0138 1.0000
6.750 0.7921 0.02927 0.02431 -0.0140 0.0162 1.0000
9.250 0.7435 0.05884 0.05660 0.0052 0.0165 1.0000
9.500 0.6833 0.07510 0.07307 -0.0074 0.0177 1.0000
9.750 0.6668 0.08391 0.08185 -0.0125 0.0180 1.0000
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