RAF 15 AIRFOIL (raf15-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 15 AIRFOIL (raf15-il) Reynolds number: 500,000 Max Cl/Cd: 78.87 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf15-il-500000.txt Download as CSV file: xf-raf15-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 15 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5273 0.08790 0.08564 -0.0059 1.0000 0.0198
-8.000 -0.5338 0.08471 0.08248 -0.0067 1.0000 0.0202
-7.750 -0.5415 0.08151 0.07930 -0.0070 1.0000 0.0209
-7.500 -0.5415 0.07744 0.07523 -0.0089 1.0000 0.0213
-7.250 -0.5378 0.07169 0.06944 -0.0135 1.0000 0.0229
-7.000 -0.5308 0.06697 0.06466 -0.0155 1.0000 0.0232
-6.750 -0.5251 0.06150 0.05911 -0.0169 1.0000 0.0233
-6.500 -0.5308 0.05307 0.05055 -0.0176 1.0000 0.0236
-6.250 -0.5277 0.04879 0.04619 -0.0165 1.0000 0.0241
-6.000 -0.5173 0.04709 0.04444 -0.0150 1.0000 0.0247
-5.750 -0.5480 0.02657 0.02241 -0.0089 1.0000 0.0180
-5.500 -0.5375 0.02373 0.01916 -0.0057 1.0000 0.0179
-5.250 -0.5265 0.02089 0.01593 -0.0026 1.0000 0.0182
-5.000 -0.5110 0.01917 0.01400 -0.0003 1.0000 0.0186
-4.750 -0.4930 0.01802 0.01266 0.0018 1.0000 0.0191
-4.500 -0.4730 0.01760 0.01222 0.0033 1.0000 0.0202
-4.250 -0.4533 0.01675 0.01119 0.0052 1.0000 0.0212
-4.000 -0.4334 0.01574 0.00995 0.0071 1.0000 0.0221
-3.750 -0.4108 0.01539 0.00943 0.0085 0.9998 0.0231
-3.500 -0.3782 0.01353 0.00736 0.0077 0.9975 0.0246
-3.250 -0.3375 0.01302 0.00684 0.0049 0.9942 0.0270
-3.000 -0.3013 0.01247 0.00623 0.0034 0.9905 0.0293
-2.750 -0.2658 0.01215 0.00584 0.0019 0.9876 0.0308
-2.500 -0.2325 0.01117 0.00482 0.0009 0.9854 0.0351
-2.250 -0.1996 0.01076 0.00440 0.0000 0.9814 0.0386
-2.000 -0.1645 0.01037 0.00397 -0.0013 0.9772 0.0414
-1.750 -0.1279 0.00992 0.00355 -0.0031 0.9742 0.0490
-1.500 -0.0976 0.00952 0.00319 -0.0033 0.9677 0.0602
-1.250 -0.0637 0.00895 0.00299 -0.0045 0.9626 0.1456
-1.000 -0.0342 0.00872 0.00286 -0.0047 0.9543 0.1779
-0.750 0.0001 0.00850 0.00270 -0.0058 0.9470 0.1971
-0.500 0.0308 0.00827 0.00250 -0.0061 0.9361 0.2136
-0.250 0.0631 0.00802 0.00229 -0.0067 0.9240 0.2279
0.000 0.1062 0.00769 0.00200 -0.0098 0.9111 0.2410
0.250 0.1775 0.00733 0.00168 -0.0193 0.8904 0.2659
0.500 0.2578 0.00724 0.00151 -0.0312 0.8200 0.2968
0.750 0.2815 0.00750 0.00145 -0.0299 0.7134 0.3260
1.000 0.2844 0.00684 0.00140 -0.0245 0.6435 0.6371
1.250 0.4551 0.00772 0.00240 -0.0568 0.4769 0.9915
1.500 0.5040 0.00798 0.00239 -0.0615 0.4249 0.9980
1.750 0.5370 0.00814 0.00239 -0.0627 0.3931 1.0000
2.000 0.5587 0.00831 0.00245 -0.0613 0.3707 1.0000
2.250 0.5807 0.00847 0.00252 -0.0601 0.3523 1.0000
2.500 0.6032 0.00862 0.00259 -0.0588 0.3360 1.0000
2.750 0.6258 0.00877 0.00269 -0.0576 0.3203 1.0000
3.000 0.6485 0.00892 0.00278 -0.0565 0.3042 1.0000
3.250 0.6714 0.00906 0.00287 -0.0553 0.2889 1.0000
3.500 0.6944 0.00921 0.00297 -0.0542 0.2731 1.0000
4.000 0.7402 0.00955 0.00320 -0.0520 0.2371 1.0000
4.250 0.7630 0.00974 0.00334 -0.0508 0.2181 1.0000
4.500 0.7855 0.00996 0.00349 -0.0497 0.1973 1.0000
4.750 0.8075 0.01024 0.00369 -0.0484 0.1753 1.0000
5.000 0.8295 0.01054 0.00391 -0.0472 0.1542 1.0000
5.250 0.8511 0.01088 0.00417 -0.0459 0.1322 1.0000
5.500 0.8718 0.01133 0.00448 -0.0445 0.1020 1.0000
5.750 0.8909 0.01197 0.00492 -0.0428 0.0704 1.0000
6.000 0.9116 0.01245 0.00536 -0.0412 0.0605 1.0000
6.250 0.9323 0.01291 0.00586 -0.0397 0.0556 1.0000
6.500 0.9535 0.01331 0.00631 -0.0383 0.0523 1.0000
6.750 0.9735 0.01383 0.00684 -0.0367 0.0490 1.0000
7.000 0.9916 0.01453 0.00761 -0.0347 0.0459 1.0000
7.250 1.0129 0.01489 0.00805 -0.0334 0.0444 1.0000
7.500 1.0333 0.01535 0.00858 -0.0319 0.0425 1.0000
7.750 1.0535 0.01580 0.00907 -0.0304 0.0402 1.0000
8.000 1.0708 0.01654 0.00987 -0.0284 0.0379 1.0000
8.250 1.0863 0.01750 0.01092 -0.0261 0.0357 1.0000
8.500 1.1072 0.01786 0.01137 -0.0248 0.0341 1.0000
8.750 1.1267 0.01835 0.01194 -0.0233 0.0322 1.0000
9.000 1.1458 0.01885 0.01247 -0.0218 0.0300 1.0000
9.250 1.1553 0.02041 0.01412 -0.0187 0.0275 1.0000
9.500 1.1771 0.02057 0.01439 -0.0176 0.0260 1.0000
9.750 1.1961 0.02101 0.01492 -0.0160 0.0241 1.0000
10.000 1.2139 0.02153 0.01547 -0.0144 0.0224 1.0000
10.250 1.2206 0.02322 0.01725 -0.0110 0.0207 1.0000
10.500 1.2385 0.02368 0.01785 -0.0094 0.0197 1.0000
10.750 1.2564 0.02409 0.01833 -0.0078 0.0182 1.0000
11.000 1.2721 0.02465 0.01896 -0.0059 0.0172 1.0000
11.250 1.2817 0.02570 0.02006 -0.0032 0.0161 1.0000
11.500 1.2841 0.02747 0.02199 0.0005 0.0155 1.0000
11.750 1.2949 0.02836 0.02303 0.0030 0.0149 1.0000
12.000 1.2990 0.02940 0.02422 0.0066 0.0143 1.0000
12.250 1.3007 0.03049 0.02542 0.0105 0.0138 1.0000
12.500 1.3008 0.03176 0.02682 0.0141 0.0134 1.0000
12.750 1.3011 0.03312 0.02830 0.0174 0.0131 1.0000
13.000 1.3019 0.03449 0.02976 0.0201 0.0127 1.0000
13.250 1.2964 0.03655 0.03194 0.0227 0.0124 1.0000
13.500 1.2887 0.03910 0.03466 0.0245 0.0122 1.0000
13.750 1.2750 0.04269 0.03841 0.0253 0.0120 1.0000
14.000 1.2554 0.04778 0.04372 0.0243 0.0118 1.0000
14.250 1.2350 0.05398 0.05014 0.0214 0.0117 1.0000
14.500 1.2294 0.05864 0.05495 0.0184 0.0116 1.0000
14.750 1.2109 0.06572 0.06223 0.0141 0.0115 1.0000
15.000 1.1870 0.07356 0.07023 0.0098 0.0117 1.0000
15.250 1.1684 0.08093 0.07773 0.0055 0.0116 1.0000
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Polar data table (+)
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