RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 25 AIRFOIL (raf25-il) Reynolds number: 500,000 Max Cl/Cd: 55.7 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf25-il-500000-n5.txt Download as CSV file: xf-raf25-il-500000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5947 0.08387 0.08165 -0.0106 1.0000 0.0050
-8.750 -0.5993 0.07902 0.07684 -0.0134 1.0000 0.0048
-8.500 -0.6068 0.07404 0.07189 -0.0166 1.0000 0.0048
-8.250 -0.6173 0.06751 0.06539 -0.0227 1.0000 0.0047
-8.000 -0.6217 0.05991 0.05772 -0.0288 1.0000 0.0046
-7.750 -0.6263 0.05179 0.04944 -0.0324 1.0000 0.0045
-7.500 -0.6340 0.04187 0.03917 -0.0338 1.0000 0.0043
-7.250 -0.6565 0.02569 0.02182 -0.0308 1.0000 0.0041
-7.000 -0.6441 0.02189 0.01749 -0.0289 1.0000 0.0042
-6.750 -0.6268 0.01956 0.01478 -0.0274 1.0000 0.0044
-6.500 -0.6078 0.01770 0.01258 -0.0259 1.0000 0.0047
-6.250 -0.5877 0.01619 0.01079 -0.0246 1.0000 0.0050
-6.000 -0.5668 0.01493 0.00930 -0.0233 1.0000 0.0053
-5.750 -0.5449 0.01408 0.00827 -0.0222 1.0000 0.0057
-5.500 -0.5241 0.01299 0.00704 -0.0209 1.0000 0.0065
-5.250 -0.5015 0.01251 0.00649 -0.0200 1.0000 0.0072
-5.000 -0.4793 0.01187 0.00573 -0.0189 1.0000 0.0080
-4.750 -0.4571 0.01127 0.00501 -0.0178 1.0000 0.0088
-4.500 -0.4260 0.01052 0.00413 -0.0186 0.9978 0.0108
-4.250 -0.3931 0.01011 0.00372 -0.0198 0.9951 0.0148
-4.000 -0.3611 0.00975 0.00335 -0.0209 0.9917 0.0212
-3.750 -0.3274 0.00965 0.00319 -0.0222 0.9882 0.0275
-3.500 -0.2950 0.00932 0.00283 -0.0234 0.9834 0.0322
-3.250 -0.2616 0.00909 0.00257 -0.0247 0.9783 0.0360
-3.000 -0.2306 0.00888 0.00231 -0.0254 0.9713 0.0385
-2.500 -0.1666 0.00838 0.00177 -0.0273 0.9581 0.0491
-2.250 -0.1329 0.00804 0.00156 -0.0287 0.9515 0.0842
-2.000 -0.1002 0.00765 0.00137 -0.0300 0.9426 0.1488
-1.750 -0.0667 0.00723 0.00121 -0.0314 0.9332 0.2239
-1.500 -0.0333 0.00687 0.00109 -0.0329 0.9226 0.3038
-1.250 -0.0025 0.00645 0.00098 -0.0337 0.9098 0.4036
-1.000 0.0262 0.00615 0.00093 -0.0340 0.8952 0.4917
-0.750 0.0533 0.00597 0.00089 -0.0338 0.8765 0.5528
-0.500 0.0784 0.00577 0.00086 -0.0331 0.8546 0.6232
-0.250 0.1015 0.00554 0.00086 -0.0319 0.8346 0.7063
0.000 0.1228 0.00521 0.00089 -0.0302 0.8190 0.8150
0.250 0.1599 0.00506 0.00094 -0.0317 0.7915 0.9214
0.500 0.2080 0.00523 0.00093 -0.0361 0.7385 0.9612
0.750 0.2428 0.00543 0.00094 -0.0375 0.6897 0.9790
1.000 0.2745 0.00568 0.00097 -0.0384 0.6323 0.9910
1.250 0.3078 0.00607 0.00101 -0.0398 0.5472 1.0000
1.500 0.3298 0.00641 0.00108 -0.0386 0.4815 1.0000
1.750 0.3515 0.00682 0.00120 -0.0375 0.3999 1.0000
2.000 0.3731 0.00732 0.00135 -0.0364 0.3123 1.0000
2.250 0.3948 0.00784 0.00153 -0.0354 0.2213 1.0000
2.500 0.4158 0.00849 0.00179 -0.0343 0.1219 1.0000
2.750 0.4383 0.00899 0.00204 -0.0333 0.0615 1.0000
3.000 0.4628 0.00921 0.00227 -0.0326 0.0537 1.0000
3.250 0.4872 0.00945 0.00252 -0.0320 0.0475 1.0000
3.500 0.5116 0.00970 0.00278 -0.0313 0.0411 1.0000
3.750 0.5362 0.00992 0.00302 -0.0306 0.0348 1.0000
4.000 0.5609 0.01015 0.00326 -0.0300 0.0269 1.0000
4.250 0.5848 0.01050 0.00354 -0.0292 0.0162 1.0000
4.500 0.6083 0.01096 0.00406 -0.0283 0.0117 1.0000
4.750 0.6319 0.01141 0.00461 -0.0274 0.0100 1.0000
5.000 0.6552 0.01188 0.00513 -0.0265 0.0084 1.0000
5.250 0.6767 0.01270 0.00605 -0.0253 0.0070 1.0000
5.500 0.6993 0.01338 0.00684 -0.0242 0.0065 1.0000
5.750 0.7214 0.01417 0.00774 -0.0231 0.0060 1.0000
6.000 0.7433 0.01508 0.00877 -0.0219 0.0055 1.0000
6.250 0.7652 0.01605 0.00990 -0.0208 0.0052 1.0000
6.500 0.7871 0.01705 0.01102 -0.0198 0.0049 1.0000
6.750 0.8065 0.01872 0.01287 -0.0185 0.0044 1.0000
7.000 0.8291 0.01967 0.01402 -0.0176 0.0041 1.0000
7.250 0.8488 0.02157 0.01623 -0.0162 0.0038 1.0000
7.500 0.8658 0.02418 0.01925 -0.0144 0.0036 1.0000
7.750 0.8787 0.02782 0.02339 -0.0122 0.0034 1.0000
8.000 0.8801 0.03466 0.03091 -0.0086 0.0032 1.0000
8.250 0.8709 0.04357 0.04042 -0.0048 0.0033 1.0000
8.500 0.8639 0.05038 0.04759 -0.0025 0.0033 1.0000
8.750 0.8538 0.05659 0.05406 -0.0010 0.0033 1.0000
9.000 0.8419 0.06193 0.05958 -0.0002 0.0034 1.0000
9.250 0.8218 0.06676 0.06452 0.0007 0.0034 1.0000
9.500 0.8016 0.07235 0.07020 -0.0018 0.0035 1.0000
9.750 0.7847 0.08156 0.07944 -0.0107 0.0035 1.0000
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Polar data table (+)
Polar graphs
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