RAF 27 AIRFOIL (raf27-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 27 AIRFOIL (raf27-il) Reynolds number: 200,000 Max Cl/Cd: 49.1 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf27-il-200000.txt Download as CSV file: xf-raf27-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 27 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.5094 0.12403 0.12047 -0.0025 1.0000 0.0635
-12.000 -0.6558 0.11795 0.11424 -0.0073 1.0000 0.0578
-11.750 -0.6317 0.11853 0.11476 -0.0022 1.0000 0.0605
-11.500 -0.6283 0.11503 0.11127 -0.0031 1.0000 0.0629
-11.250 -0.6313 0.11021 0.10647 -0.0055 1.0000 0.0653
-9.750 -0.8527 0.05266 0.04799 -0.0293 1.0000 0.0368
-9.500 -0.8512 0.05144 0.04639 -0.0261 1.0000 0.0356
-9.250 -0.8568 0.04718 0.04181 -0.0234 1.0000 0.0357
-9.000 -0.8511 0.04306 0.03749 -0.0215 1.0000 0.0349
-8.750 -0.8488 0.03933 0.03341 -0.0187 1.0000 0.0347
-8.500 -0.8434 0.03577 0.02947 -0.0160 1.0000 0.0346
-8.250 -0.8338 0.03287 0.02616 -0.0134 1.0000 0.0347
-8.000 -0.8206 0.03074 0.02365 -0.0110 1.0000 0.0351
-7.750 -0.8071 0.02761 0.02016 -0.0089 1.0000 0.0360
-7.500 -0.7897 0.02531 0.01779 -0.0076 1.0000 0.0383
-7.250 -0.7704 0.02405 0.01642 -0.0061 1.0000 0.0406
-7.000 -0.7501 0.02258 0.01477 -0.0046 1.0000 0.0426
-6.750 -0.7292 0.02138 0.01337 -0.0031 1.0000 0.0447
-6.500 -0.7103 0.01970 0.01166 -0.0016 1.0000 0.0483
-6.250 -0.6908 0.01882 0.01077 0.0000 1.0000 0.0523
-6.000 -0.6708 0.01798 0.00984 0.0016 1.0000 0.0560
-5.750 -0.6539 0.01685 0.00875 0.0035 1.0000 0.0611
-5.500 -0.6346 0.01623 0.00809 0.0052 1.0000 0.0678
-5.250 -0.6180 0.01532 0.00722 0.0073 1.0000 0.0752
-5.000 -0.6007 0.01456 0.00649 0.0092 1.0000 0.0872
-4.750 -0.5849 0.01363 0.00571 0.0114 1.0000 0.1133
-4.500 -0.5733 0.01232 0.00507 0.0140 1.0000 0.2189
-4.250 -0.5579 0.01159 0.00472 0.0160 1.0000 0.3081
-4.000 -0.5413 0.01108 0.00449 0.0180 1.0000 0.3737
-3.750 -0.5243 0.01068 0.00435 0.0200 1.0000 0.4383
-3.500 -0.5070 0.01037 0.00424 0.0220 1.0000 0.4953
-3.250 -0.4897 0.01011 0.00417 0.0241 1.0000 0.5470
-3.000 -0.4730 0.00989 0.00416 0.0263 1.0000 0.6010
-2.750 -0.4567 0.00975 0.00422 0.0287 1.0000 0.6546
-2.500 -0.4398 0.00966 0.00427 0.0310 1.0000 0.6982
-2.250 -0.4228 0.00960 0.00432 0.0331 1.0000 0.7365
-2.000 -0.3955 0.00960 0.00444 0.0333 0.9970 0.7795
-1.750 -0.3558 0.00966 0.00462 0.0312 0.9904 0.8217
-1.500 -0.3131 0.00976 0.00478 0.0284 0.9845 0.8563
-1.250 -0.2699 0.00983 0.00486 0.0254 0.9781 0.8838
-1.000 -0.2198 0.00997 0.00499 0.0211 0.9735 0.9050
-0.750 -0.1650 0.01013 0.00512 0.0158 0.9702 0.9216
-0.500 -0.1139 0.01024 0.00521 0.0112 0.9634 0.9349
-0.250 -0.0570 0.01031 0.00526 0.0056 0.9581 0.9446
0.000 0.0001 0.01034 0.00529 0.0000 0.9516 0.9517
0.250 0.0570 0.01031 0.00526 -0.0056 0.9446 0.9581
0.500 0.1139 0.01023 0.00520 -0.0112 0.9348 0.9633
0.750 0.1650 0.01013 0.00512 -0.0158 0.9214 0.9702
1.000 0.2197 0.00997 0.00499 -0.0211 0.9050 0.9735
1.250 0.2699 0.00983 0.00486 -0.0254 0.8839 0.9782
1.500 0.3132 0.00976 0.00478 -0.0284 0.8565 0.9845
1.750 0.3558 0.00966 0.00462 -0.0311 0.8214 0.9904
2.000 0.3954 0.00960 0.00444 -0.0333 0.7787 0.9970
2.250 0.4228 0.00960 0.00432 -0.0331 0.7364 1.0000
2.500 0.4400 0.00966 0.00427 -0.0310 0.6989 1.0000
2.750 0.4569 0.00975 0.00422 -0.0288 0.6556 1.0000
3.000 0.4731 0.00989 0.00416 -0.0264 0.6013 1.0000
3.250 0.4898 0.01011 0.00418 -0.0241 0.5470 1.0000
3.500 0.5071 0.01037 0.00424 -0.0221 0.4955 1.0000
3.750 0.5244 0.01068 0.00435 -0.0200 0.4384 1.0000
4.000 0.5414 0.01108 0.00449 -0.0180 0.3736 1.0000
4.250 0.5581 0.01159 0.00472 -0.0161 0.3083 1.0000
4.500 0.5733 0.01233 0.00507 -0.0140 0.2179 1.0000
4.750 0.5851 0.01363 0.00571 -0.0114 0.1133 1.0000
5.000 0.6007 0.01458 0.00650 -0.0092 0.0871 1.0000
5.250 0.6181 0.01532 0.00722 -0.0073 0.0752 1.0000
5.500 0.6347 0.01624 0.00809 -0.0052 0.0678 1.0000
5.750 0.6540 0.01685 0.00876 -0.0036 0.0613 1.0000
6.000 0.6709 0.01798 0.00984 -0.0016 0.0560 1.0000
6.250 0.6909 0.01881 0.01077 0.0000 0.0523 1.0000
6.500 0.7104 0.01971 0.01166 0.0015 0.0483 1.0000
6.750 0.7293 0.02141 0.01340 0.0031 0.0447 1.0000
7.000 0.7502 0.02260 0.01479 0.0046 0.0426 1.0000
7.250 0.7705 0.02407 0.01645 0.0061 0.0407 1.0000
7.500 0.7898 0.02538 0.01787 0.0076 0.0385 1.0000
7.750 0.8075 0.02742 0.01997 0.0089 0.0362 1.0000
8.000 0.8206 0.03089 0.02380 0.0110 0.0352 1.0000
8.250 0.8338 0.03289 0.02618 0.0134 0.0347 1.0000
8.500 0.8433 0.03575 0.02946 0.0160 0.0345 1.0000
8.750 0.8489 0.03925 0.03334 0.0187 0.0347 1.0000
9.000 0.8513 0.04458 0.03893 0.0210 0.0355 1.0000
9.250 0.8494 0.04818 0.04287 0.0238 0.0353 1.0000
9.500 0.8453 0.05164 0.04664 0.0266 0.0353 1.0000
9.750 0.8496 0.05289 0.04827 0.0296 0.0369 1.0000
10.000 0.5759 0.07532 0.07194 0.0221 0.0768 1.0000
10.250 0.5573 0.08295 0.07954 0.0182 0.0760 1.0000
10.500 0.5461 0.08911 0.08567 0.0157 0.0749 1.0000
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