RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 25 AIRFOIL (raf25-il) Reynolds number: 200,000 Max Cl/Cd: 56.2 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf25-il-200000.txt Download as CSV file: xf-raf25-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5743 0.09440 0.09084 -0.0155 1.0000 0.0399
-8.750 -0.5764 0.08992 0.08640 -0.0193 1.0000 0.0400
-8.500 -0.5828 0.08474 0.08125 -0.0248 1.0000 0.0401
-8.250 -0.5837 0.08034 0.07693 -0.0192 1.0000 0.0421
-8.000 -0.5812 0.07734 0.07394 -0.0193 1.0000 0.0433
-7.750 -0.5800 0.07294 0.06954 -0.0225 1.0000 0.0442
-7.500 -0.5776 0.06840 0.06497 -0.0255 1.0000 0.0455
-7.250 -0.5745 0.06351 0.06001 -0.0285 1.0000 0.0473
-7.000 -0.5695 0.05837 0.05470 -0.0313 1.0000 0.0497
-6.750 -0.5634 0.05549 0.05107 -0.0334 1.0000 0.0527
-6.500 -0.5589 0.04793 0.04368 -0.0337 1.0000 0.0547
-6.250 -0.5452 0.04519 0.04095 -0.0330 1.0000 0.0570
-6.000 -0.5282 0.04524 0.04023 -0.0318 1.0000 0.0657
-5.750 -0.4912 0.02428 0.01988 -0.0323 1.0000 0.0695
-5.500 -0.4760 0.02176 0.01717 -0.0313 1.0000 0.0756
-5.250 -0.4628 0.01875 0.01394 -0.0304 1.0000 0.0840
-5.000 -0.4608 0.02551 0.01871 -0.0255 1.0000 0.0404
-4.750 -0.4405 0.02213 0.01499 -0.0240 1.0000 0.0393
-4.500 -0.4183 0.02019 0.01268 -0.0226 1.0000 0.0398
-4.250 -0.3968 0.01741 0.00963 -0.0212 1.0000 0.0422
-4.000 -0.3746 0.01638 0.00857 -0.0202 1.0000 0.0476
-3.750 -0.3516 0.01527 0.00730 -0.0189 1.0000 0.0509
-3.500 -0.3300 0.01400 0.00599 -0.0176 1.0000 0.0557
-3.250 -0.3081 0.01339 0.00538 -0.0164 1.0000 0.0633
-3.000 -0.2870 0.01255 0.00454 -0.0151 1.0000 0.0690
-2.750 -0.2652 0.01209 0.00406 -0.0140 1.0000 0.0794
-2.500 -0.2442 0.01142 0.00347 -0.0127 1.0000 0.0966
-2.250 -0.2273 0.00974 0.00305 -0.0114 1.0000 0.3660
-2.000 -0.2125 0.00890 0.00310 -0.0090 1.0000 0.5843
-1.750 -0.1978 0.00838 0.00315 -0.0060 1.0000 0.7270
-1.500 -0.0999 0.00789 0.00303 -0.0193 1.0000 1.0000
-1.250 -0.0841 0.00796 0.00299 -0.0173 1.0000 1.0000
-1.000 -0.0677 0.00806 0.00299 -0.0154 1.0000 1.0000
-0.750 -0.0508 0.00819 0.00304 -0.0137 1.0000 1.0000
-0.500 -0.0335 0.00835 0.00312 -0.0121 1.0000 1.0000
-0.250 -0.0016 0.00852 0.00322 -0.0135 0.9968 1.0000
0.000 0.0430 0.00866 0.00331 -0.0174 0.9898 1.0000
0.250 0.0906 0.00878 0.00341 -0.0218 0.9831 1.0000
0.500 0.1368 0.00878 0.00341 -0.0258 0.9731 1.0000
0.750 0.1827 0.00874 0.00340 -0.0296 0.9637 1.0000
1.000 0.2307 0.00868 0.00339 -0.0338 0.9569 1.0000
1.250 0.2770 0.00852 0.00331 -0.0375 0.9456 1.0000
1.500 0.3289 0.00812 0.00298 -0.0416 0.9217 1.0000
1.750 0.3701 0.00791 0.00280 -0.0435 0.8921 1.0000
2.000 0.4007 0.00788 0.00277 -0.0434 0.8610 1.0000
2.250 0.4249 0.00794 0.00278 -0.0417 0.8183 1.0000
2.500 0.4465 0.00808 0.00279 -0.0396 0.7641 1.0000
2.750 0.4670 0.00831 0.00283 -0.0374 0.6938 1.0000
3.000 0.4837 0.00885 0.00288 -0.0344 0.5714 1.0000
3.250 0.4899 0.01063 0.00321 -0.0304 0.2663 1.0000
3.500 0.5043 0.01237 0.00402 -0.0283 0.0892 1.0000
3.750 0.5260 0.01314 0.00472 -0.0271 0.0699 1.0000
4.000 0.5457 0.01428 0.00584 -0.0255 0.0593 1.0000
4.250 0.5670 0.01529 0.00685 -0.0242 0.0494 1.0000
4.500 0.5883 0.01687 0.00846 -0.0227 0.0449 1.0000
4.750 0.6125 0.01813 0.00986 -0.0216 0.0413 1.0000
5.000 0.6358 0.01943 0.01114 -0.0208 0.0361 1.0000
5.250 0.6596 0.02303 0.01495 -0.0198 0.0343 1.0000
5.500 0.6833 0.02506 0.01728 -0.0186 0.0338 1.0000
5.750 0.7068 0.02776 0.02026 -0.0173 0.0347 1.0000
6.000 0.7278 0.03052 0.02337 -0.0159 0.0348 1.0000
6.250 0.7471 0.03360 0.02681 -0.0142 0.0348 1.0000
7.250 0.8054 0.04918 0.04441 -0.0054 0.0495 1.0000
7.500 0.8182 0.05387 0.04903 -0.0051 0.0480 1.0000
7.750 0.8126 0.06240 0.05778 -0.0046 0.0468 1.0000
8.000 0.8140 0.06478 0.06055 -0.0030 0.0464 1.0000
8.250 0.8072 0.06688 0.06316 -0.0018 0.0441 1.0000
8.500 0.7986 0.07144 0.06790 -0.0021 0.0429 1.0000
8.750 0.7859 0.07581 0.07236 -0.0025 0.0425 1.0000
9.000 0.7694 0.08074 0.07735 -0.0047 0.0424 1.0000
9.250 0.7553 0.08759 0.08419 -0.0110 0.0426 1.0000
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Polar data table (+)
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