RAF 15 AIRFOIL (raf15-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 15 AIRFOIL (raf15-il) Reynolds number: 200,000 Max Cl/Cd: 59.46 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf15-il-200000.txt Download as CSV file: xf-raf15-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5145 0.08907 0.08554 -0.0082 1.0000 0.0442
-7.750 -0.5183 0.08623 0.08273 -0.0083 1.0000 0.0453
-7.500 -0.5202 0.08302 0.07956 -0.0090 1.0000 0.0463
-7.250 -0.5194 0.07944 0.07598 -0.0105 1.0000 0.0480
-6.750 -0.5096 0.07050 0.06678 -0.0201 1.0000 0.0521
-6.500 -0.5118 0.06533 0.06171 -0.0179 1.0000 0.0533
-6.250 -0.5035 0.06312 0.05953 -0.0157 1.0000 0.0547
-6.000 -0.4936 0.06080 0.05720 -0.0144 1.0000 0.0575
-5.750 -0.4811 0.05481 0.05072 -0.0181 1.0000 0.0647
-5.500 -0.4738 0.05249 0.04861 -0.0154 1.0000 0.0666
-5.250 -0.4615 0.05091 0.04704 -0.0134 1.0000 0.0702
-5.000 -0.4495 0.04601 0.04176 -0.0136 1.0000 0.0787
-4.750 -0.4369 0.04402 0.03982 -0.0117 1.0000 0.0806
-4.500 -0.4232 0.04096 0.03629 -0.0102 1.0000 0.0916
-4.250 -0.4259 0.02461 0.01833 -0.0034 1.0000 0.0429
-4.000 -0.4078 0.02237 0.01552 -0.0007 1.0000 0.0416
-3.750 -0.3897 0.02072 0.01379 0.0010 1.0000 0.0448
-3.500 -0.3692 0.01912 0.01187 0.0030 1.0000 0.0456
-3.250 -0.3471 0.01773 0.01019 0.0047 1.0000 0.0465
-3.000 -0.3244 0.01664 0.00890 0.0063 1.0000 0.0483
-2.750 -0.3019 0.01606 0.00813 0.0078 1.0000 0.0516
-2.500 -0.2796 0.01484 0.00687 0.0092 1.0000 0.0550
-2.250 -0.2582 0.01424 0.00628 0.0107 1.0000 0.0587
-2.000 -0.2367 0.01388 0.00586 0.0123 1.0000 0.0641
-1.750 -0.2165 0.01330 0.00536 0.0140 1.0000 0.0717
-1.500 -0.1961 0.01293 0.00504 0.0156 1.0000 0.0832
-1.250 -0.1765 0.01241 0.00482 0.0174 1.0000 0.1290
-1.000 -0.1557 0.01247 0.00502 0.0188 1.0000 0.1924
-0.750 -0.1243 0.01267 0.00519 0.0178 0.9973 0.2205
-0.500 -0.0782 0.01274 0.00533 0.0138 0.9897 0.2463
-0.250 -0.0324 0.01264 0.00533 0.0100 0.9815 0.2716
0.000 0.0144 0.01240 0.00521 0.0059 0.9731 0.3002
0.250 0.0554 0.01195 0.00492 0.0033 0.9620 0.3275
0.500 0.0946 0.01136 0.00454 0.0011 0.9493 0.3622
0.750 0.2500 0.00999 0.00494 -0.0264 0.9653 1.0000
1.000 0.2916 0.00962 0.00454 -0.0288 0.9375 1.0000
1.250 0.3593 0.00908 0.00395 -0.0365 0.9010 1.0000
1.500 0.5103 0.00913 0.00315 -0.0623 0.6792 1.0000
1.750 0.5266 0.00972 0.00320 -0.0594 0.5833 1.0000
2.000 0.5441 0.01018 0.00331 -0.0570 0.5226 1.0000
2.250 0.5636 0.01055 0.00344 -0.0552 0.4802 1.0000
2.500 0.5844 0.01087 0.00358 -0.0535 0.4492 1.0000
2.750 0.6057 0.01116 0.00374 -0.0521 0.4240 1.0000
3.000 0.6275 0.01145 0.00394 -0.0507 0.4026 1.0000
3.250 0.6493 0.01176 0.00413 -0.0494 0.3838 1.0000
3.500 0.6715 0.01202 0.00435 -0.0481 0.3649 1.0000
3.750 0.6931 0.01229 0.00455 -0.0467 0.3435 1.0000
4.000 0.7142 0.01256 0.00475 -0.0453 0.3206 1.0000
4.250 0.7360 0.01279 0.00496 -0.0439 0.2997 1.0000
4.500 0.7578 0.01306 0.00520 -0.0426 0.2816 1.0000
4.750 0.7796 0.01332 0.00544 -0.0413 0.2625 1.0000
5.000 0.8015 0.01355 0.00571 -0.0400 0.2407 1.0000
5.250 0.8229 0.01384 0.00595 -0.0386 0.2172 1.0000
5.500 0.8438 0.01420 0.00624 -0.0372 0.1901 1.0000
5.750 0.8639 0.01470 0.00664 -0.0357 0.1584 1.0000
6.000 0.8815 0.01552 0.00727 -0.0337 0.1228 1.0000
6.250 0.8993 0.01639 0.00804 -0.0317 0.0995 1.0000
6.500 0.9174 0.01722 0.00885 -0.0297 0.0880 1.0000
6.750 0.9349 0.01815 0.00972 -0.0277 0.0808 1.0000
7.000 0.9545 0.01888 0.01055 -0.0260 0.0759 1.0000
7.250 0.9728 0.01979 0.01148 -0.0242 0.0716 1.0000
7.500 0.9911 0.02097 0.01270 -0.0224 0.0678 1.0000
7.750 1.0108 0.02179 0.01367 -0.0208 0.0643 1.0000
8.000 1.0297 0.02272 0.01463 -0.0192 0.0607 1.0000
8.250 1.0471 0.02459 0.01653 -0.0176 0.0571 1.0000
8.500 1.0657 0.02552 0.01773 -0.0158 0.0545 1.0000
8.750 1.0830 0.02653 0.01890 -0.0139 0.0511 1.0000
9.000 1.0995 0.02793 0.02031 -0.0123 0.0478 1.0000
9.250 1.1122 0.03050 0.02317 -0.0100 0.0453 1.0000
9.500 1.1254 0.03174 0.02472 -0.0075 0.0426 1.0000
9.750 1.1382 0.03305 0.02622 -0.0053 0.0400 1.0000
10.000 1.1499 0.03479 0.02808 -0.0031 0.0380 1.0000
10.250 1.1515 0.03913 0.03269 -0.0003 0.0363 1.0000
10.500 1.1518 0.04130 0.03525 0.0034 0.0354 1.0000
10.750 1.1499 0.04346 0.03779 0.0072 0.0343 1.0000
11.000 1.1452 0.04578 0.04043 0.0109 0.0331 1.0000
11.250 1.1371 0.04811 0.04301 0.0147 0.0320 1.0000
11.500 1.1179 0.05109 0.04625 0.0194 0.0320 1.0000
11.750 1.0995 0.05413 0.04951 0.0226 0.0316 1.0000
12.000 1.0764 0.05818 0.05379 0.0239 0.0318 1.0000
12.250 1.0491 0.06369 0.05953 0.0224 0.0320 1.0000
12.500 1.0185 0.07139 0.06745 0.0172 0.0327 1.0000
12.750 0.9831 0.08199 0.07820 0.0093 0.0342 1.0000
13.000 0.9516 0.09274 0.08901 0.0022 0.0354 1.0000
13.250 0.9178 0.10465 0.10093 -0.0047 0.0360 1.0000
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Polar data table (+)
Polar graphs
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