RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 25 AIRFOIL (raf25-il) Reynolds number: 100,000 Max Cl/Cd: 43.28 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf25-il-100000.txt Download as CSV file: xf-raf25-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5670 0.09018 0.08530 -0.0097 1.0000 0.1016
-8.000 -0.5820 0.08650 0.08173 -0.0155 1.0000 0.1047
-7.750 -0.6032 0.08195 0.07705 -0.0271 1.0000 0.1061
-7.500 -0.5726 0.07861 0.07390 -0.0162 1.0000 0.1129
-7.250 -0.5901 0.07380 0.06883 -0.0283 1.0000 0.1194
-7.000 -0.5680 0.07050 0.06575 -0.0210 1.0000 0.1277
-6.750 -0.5651 0.06620 0.06142 -0.0233 1.0000 0.1375
-6.500 -0.5620 0.06197 0.05705 -0.0265 1.0000 0.1490
-6.250 -0.5537 0.05849 0.05347 -0.0271 1.0000 0.1625
-6.000 -0.5430 0.05513 0.05005 -0.0268 1.0000 0.1765
-5.750 -0.5307 0.05190 0.04682 -0.0258 1.0000 0.1912
-5.500 -0.5173 0.04898 0.04390 -0.0243 1.0000 0.2074
-5.250 -0.4841 0.03377 0.02649 -0.0298 1.0000 0.0759
-5.000 -0.4632 0.02953 0.02125 -0.0278 1.0000 0.0668
-4.750 -0.4429 0.02706 0.01876 -0.0269 1.0000 0.0715
-4.500 -0.4202 0.02467 0.01592 -0.0255 1.0000 0.0729
-4.250 -0.3963 0.02244 0.01328 -0.0241 1.0000 0.0740
-4.000 -0.3719 0.02100 0.01142 -0.0228 1.0000 0.0788
-3.750 -0.3490 0.01914 0.00958 -0.0218 1.0000 0.0866
-3.500 -0.3247 0.01792 0.00816 -0.0205 1.0000 0.0927
-3.250 -0.3026 0.01676 0.00709 -0.0193 1.0000 0.1060
-3.000 -0.2812 0.01565 0.00610 -0.0179 1.0000 0.1190
-2.750 -0.2604 0.01458 0.00521 -0.0165 1.0000 0.1473
-2.500 -0.2499 0.01164 0.00467 -0.0133 1.0000 0.5915
-2.250 -0.2212 0.01080 0.00460 -0.0111 1.0000 0.8571
-2.000 -0.1281 0.01073 0.00413 -0.0239 1.0000 1.0000
-1.750 -0.1114 0.01069 0.00389 -0.0222 1.0000 1.0000
-1.500 -0.0939 0.01068 0.00368 -0.0205 1.0000 1.0000
-1.250 -0.0759 0.01070 0.00356 -0.0189 1.0000 1.0000
-1.000 -0.0577 0.01076 0.00348 -0.0173 1.0000 1.0000
-0.750 -0.0392 0.01084 0.00345 -0.0158 1.0000 1.0000
-0.500 -0.0207 0.01095 0.00347 -0.0144 1.0000 1.0000
-0.250 -0.0021 0.01108 0.00351 -0.0131 1.0000 1.0000
0.000 0.0165 0.01125 0.00361 -0.0118 1.0000 1.0000
0.250 0.0352 0.01144 0.00376 -0.0106 1.0000 1.0000
0.500 0.0538 0.01167 0.00394 -0.0094 1.0000 1.0000
0.750 0.0724 0.01193 0.00418 -0.0084 1.0000 1.0000
1.000 0.0909 0.01222 0.00447 -0.0074 1.0000 1.0000
1.250 0.1091 0.01255 0.00481 -0.0065 1.0000 1.0000
1.500 0.1272 0.01293 0.00521 -0.0056 1.0000 1.0000
1.750 0.1843 0.01343 0.00582 -0.0124 0.9859 1.0000
2.000 0.2387 0.01381 0.00636 -0.0183 0.9720 1.0000
2.250 0.2925 0.01405 0.00679 -0.0239 0.9576 1.0000
2.500 0.3704 0.01359 0.00666 -0.0328 0.9311 1.0000
2.750 0.4501 0.01235 0.00581 -0.0397 0.8840 1.0000
3.000 0.4868 0.01191 0.00551 -0.0390 0.8300 1.0000
3.250 0.5090 0.01176 0.00530 -0.0353 0.7393 1.0000
3.500 0.5206 0.01252 0.00506 -0.0295 0.4921 1.0000
3.750 0.5230 0.01582 0.00611 -0.0252 0.1352 1.0000
4.000 0.5425 0.01714 0.00730 -0.0234 0.1074 1.0000
4.250 0.5630 0.01836 0.00841 -0.0219 0.0886 1.0000
4.500 0.5855 0.01995 0.00997 -0.0204 0.0801 1.0000
4.750 0.6096 0.02175 0.01168 -0.0195 0.0708 1.0000
5.000 0.6361 0.02356 0.01372 -0.0185 0.0662 1.0000
5.250 0.6628 0.02585 0.01624 -0.0175 0.0644 1.0000
5.500 0.6876 0.02839 0.01898 -0.0165 0.0621 1.0000
5.750 0.7089 0.03197 0.02290 -0.0154 0.0596 1.0000
6.000 0.7329 0.03444 0.02624 -0.0128 0.0653 1.0000
6.250 0.7529 0.03986 0.03188 -0.0117 0.0724 1.0000
8.000 0.7249 0.06887 0.06452 -0.0062 0.1521 1.0000
8.250 0.6899 0.07413 0.06981 -0.0089 0.1509 1.0000
8.500 0.6608 0.08034 0.07594 -0.0140 0.1487 1.0000
8.750 0.6830 0.08381 0.07943 -0.0098 0.1387 1.0000
9.000 0.6504 0.09008 0.08562 -0.0162 0.1370 1.0000
9.250 0.6398 0.09486 0.09036 -0.0185 0.1309 1.0000
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