RAF 27 AIRFOIL (raf27-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 27 AIRFOIL (raf27-il) Reynolds number: 50,000 Max Cl/Cd: 27.55 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf27-il-50000-n5.txt Download as CSV file: xf-raf27-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6868 0.09225 0.08508 -0.0187 1.0000 0.0483
-10.750 -0.7032 0.08482 0.07766 -0.0237 1.0000 0.0478
-10.500 -0.7235 0.07806 0.07087 -0.0281 1.0000 0.0472
-10.250 -0.7449 0.07250 0.06524 -0.0308 1.0000 0.0467
-10.000 -0.7664 0.06797 0.06058 -0.0313 1.0000 0.0464
-9.750 -0.7857 0.06414 0.05660 -0.0299 1.0000 0.0464
-9.500 -0.7993 0.06025 0.05246 -0.0284 1.0000 0.0465
-9.250 -0.8085 0.05636 0.04824 -0.0265 1.0000 0.0470
-9.000 -0.8128 0.05254 0.04401 -0.0245 1.0000 0.0476
-8.750 -0.8120 0.04889 0.03989 -0.0224 1.0000 0.0485
-8.500 -0.8073 0.04555 0.03596 -0.0201 1.0000 0.0500
-8.250 -0.7956 0.04300 0.03337 -0.0188 1.0000 0.0531
-8.000 -0.7821 0.04071 0.03077 -0.0173 1.0000 0.0567
-7.750 -0.7660 0.03805 0.02768 -0.0157 1.0000 0.0599
-7.500 -0.7483 0.03575 0.02506 -0.0142 1.0000 0.0647
-7.250 -0.7300 0.03399 0.02319 -0.0130 1.0000 0.0711
-7.000 -0.7080 0.03202 0.02088 -0.0117 1.0000 0.0771
-6.750 -0.6893 0.03050 0.01936 -0.0104 1.0000 0.0863
-6.500 -0.6690 0.02890 0.01768 -0.0091 1.0000 0.0951
-6.250 -0.6492 0.02745 0.01606 -0.0075 1.0000 0.1052
-6.000 -0.6322 0.02611 0.01479 -0.0059 1.0000 0.1220
-5.750 -0.6151 0.02476 0.01351 -0.0042 1.0000 0.1438
-5.500 -0.5987 0.02340 0.01237 -0.0025 1.0000 0.1766
-5.250 -0.5836 0.02213 0.01144 -0.0007 1.0000 0.2369
-5.000 -0.5696 0.02103 0.01076 0.0014 1.0000 0.3168
-4.750 -0.5557 0.02019 0.01028 0.0039 1.0000 0.3974
-4.500 -0.5398 0.01960 0.00991 0.0064 1.0000 0.4666
-4.250 -0.5228 0.01914 0.00959 0.0089 1.0000 0.5225
-4.000 -0.5056 0.01881 0.00943 0.0117 1.0000 0.5795
-3.750 -0.4881 0.01860 0.00933 0.0146 1.0000 0.6363
-3.500 -0.4693 0.01845 0.00920 0.0172 1.0000 0.6858
-3.250 -0.4466 0.01831 0.00905 0.0191 1.0000 0.7233
-3.000 -0.4227 0.01820 0.00887 0.0206 1.0000 0.7574
-2.750 -0.3965 0.01814 0.00870 0.0216 1.0000 0.7901
-2.500 -0.3657 0.01815 0.00862 0.0217 1.0000 0.8209
-2.250 -0.3312 0.01820 0.00855 0.0210 1.0000 0.8504
-2.000 -0.2926 0.01829 0.00848 0.0193 1.0000 0.8781
-1.750 -0.2474 0.01841 0.00845 0.0162 1.0000 0.9028
-1.500 -0.1977 0.01850 0.00837 0.0120 1.0000 0.9247
-1.250 -0.1471 0.01851 0.00826 0.0074 1.0000 0.9452
-1.000 -0.0984 0.01844 0.00808 0.0028 1.0000 0.9652
-0.750 -0.0495 0.01830 0.00787 -0.0022 1.0000 0.9837
-0.500 -0.0010 0.01810 0.00763 -0.0073 1.0000 1.0000
-0.250 0.0002 0.01798 0.00753 -0.0038 1.0000 1.0000
0.000 0.0000 0.01793 0.00749 0.0000 1.0000 1.0000
0.250 -0.0003 0.01798 0.00753 0.0038 1.0000 1.0000
0.500 0.0010 0.01810 0.00763 0.0073 1.0000 1.0000
0.750 0.0493 0.01830 0.00787 0.0022 0.9837 1.0000
1.000 0.0983 0.01844 0.00808 -0.0028 0.9653 1.0000
1.250 0.1470 0.01851 0.00825 -0.0074 0.9452 1.0000
1.500 0.1977 0.01849 0.00837 -0.0120 0.9247 1.0000
1.750 0.2473 0.01841 0.00845 -0.0162 0.9028 1.0000
2.000 0.2925 0.01829 0.00848 -0.0193 0.8782 1.0000
2.250 0.3311 0.01820 0.00855 -0.0210 0.8504 1.0000
2.500 0.3657 0.01815 0.00861 -0.0217 0.8209 1.0000
2.750 0.3964 0.01814 0.00870 -0.0216 0.7901 1.0000
3.000 0.4225 0.01820 0.00886 -0.0205 0.7574 1.0000
3.250 0.4465 0.01831 0.00904 -0.0191 0.7233 1.0000
3.500 0.4692 0.01844 0.00920 -0.0172 0.6858 1.0000
3.750 0.4880 0.01860 0.00933 -0.0146 0.6363 1.0000
4.000 0.5055 0.01881 0.00943 -0.0116 0.5795 1.0000
4.250 0.5227 0.01914 0.00959 -0.0089 0.5224 1.0000
4.500 0.5397 0.01959 0.00990 -0.0064 0.4666 1.0000
4.750 0.5557 0.02018 0.01028 -0.0039 0.3979 1.0000
5.000 0.5695 0.02103 0.01075 -0.0014 0.3168 1.0000
5.250 0.5835 0.02212 0.01144 0.0007 0.2372 1.0000
5.500 0.5987 0.02340 0.01236 0.0025 0.1766 1.0000
5.750 0.6150 0.02476 0.01351 0.0042 0.1439 1.0000
6.000 0.6321 0.02611 0.01478 0.0059 0.1215 1.0000
6.500 0.6690 0.02889 0.01768 0.0091 0.0951 1.0000
6.750 0.6893 0.03050 0.01936 0.0104 0.0863 1.0000
7.000 0.7081 0.03203 0.02089 0.0117 0.0771 1.0000
7.250 0.7301 0.03400 0.02320 0.0130 0.0710 1.0000
7.500 0.7484 0.03576 0.02507 0.0142 0.0647 1.0000
7.750 0.7661 0.03805 0.02767 0.0156 0.0599 1.0000
8.000 0.7822 0.04072 0.03078 0.0173 0.0567 1.0000
8.250 0.7957 0.04302 0.03340 0.0188 0.0532 1.0000
8.500 0.8078 0.04556 0.03595 0.0201 0.0501 1.0000
8.750 0.8119 0.04893 0.03995 0.0224 0.0484 1.0000
9.000 0.8130 0.05256 0.04403 0.0245 0.0476 1.0000
9.250 0.8088 0.05637 0.04825 0.0265 0.0470 1.0000
9.500 0.7997 0.06027 0.05248 0.0283 0.0465 1.0000
9.750 0.7863 0.06414 0.05659 0.0298 0.0463 1.0000
10.000 0.7665 0.06801 0.06063 0.0312 0.0463 1.0000
10.250 0.7453 0.07254 0.06527 0.0307 0.0467 1.0000
10.500 0.7245 0.07804 0.07085 0.0280 0.0472 1.0000
10.750 0.7038 0.08489 0.07772 0.0235 0.0478 1.0000
11.000 0.6881 0.09218 0.08501 0.0187 0.0484 1.0000
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Polar data table (+)
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