RAF 27 AIRFOIL (raf27-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: RAF 27 AIRFOIL (raf27-il) Reynolds number: 100,000 Max Cl/Cd: 36.46 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf27-il-100000-n5.txt Download as CSV file: xf-raf27-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7089 0.08820 0.08311 -0.0170 1.0000 0.0240
-11.250 -0.7727 0.06978 0.06447 -0.0313 1.0000 0.0220
-11.000 -0.8030 0.06278 0.05724 -0.0353 1.0000 0.0217
-10.750 -0.8253 0.05815 0.05239 -0.0361 1.0000 0.0215
-10.500 -0.8433 0.05461 0.04864 -0.0347 1.0000 0.0216
-10.250 -0.8571 0.05175 0.04559 -0.0320 1.0000 0.0218
-10.000 -0.8655 0.04881 0.04241 -0.0295 1.0000 0.0221
-9.750 -0.8680 0.04611 0.03947 -0.0273 1.0000 0.0225
-9.500 -0.8662 0.04367 0.03677 -0.0251 1.0000 0.0233
-9.250 -0.8611 0.04141 0.03424 -0.0230 1.0000 0.0245
-9.000 -0.8546 0.03893 0.03141 -0.0208 1.0000 0.0259
-8.750 -0.8455 0.03633 0.02838 -0.0186 1.0000 0.0272
-8.500 -0.8328 0.03381 0.02544 -0.0165 1.0000 0.0283
-8.250 -0.8175 0.03161 0.02288 -0.0147 1.0000 0.0296
-8.000 -0.8023 0.02984 0.02109 -0.0133 1.0000 0.0319
-7.750 -0.7850 0.02846 0.01957 -0.0118 1.0000 0.0347
-7.500 -0.7665 0.02687 0.01771 -0.0102 1.0000 0.0375
-7.250 -0.7492 0.02535 0.01608 -0.0086 1.0000 0.0408
-7.000 -0.7317 0.02431 0.01502 -0.0071 1.0000 0.0453
-6.750 -0.7133 0.02323 0.01376 -0.0054 1.0000 0.0495
-6.500 -0.6966 0.02217 0.01273 -0.0037 1.0000 0.0552
-6.250 -0.6782 0.02133 0.01177 -0.0020 1.0000 0.0616
-6.000 -0.6610 0.02041 0.01086 -0.0002 1.0000 0.0692
-5.750 -0.6431 0.01959 0.00994 0.0015 1.0000 0.0784
-5.500 -0.6253 0.01877 0.00907 0.0033 1.0000 0.0894
-5.250 -0.6082 0.01788 0.00828 0.0051 1.0000 0.1077
-5.000 -0.5913 0.01698 0.00760 0.0068 1.0000 0.1416
-4.750 -0.5741 0.01616 0.00706 0.0085 1.0000 0.1962
-4.500 -0.5561 0.01552 0.00659 0.0101 1.0000 0.2514
-4.250 -0.5378 0.01495 0.00621 0.0117 1.0000 0.3053
-4.000 -0.5212 0.01434 0.00594 0.0137 1.0000 0.3760
-3.750 -0.5044 0.01384 0.00575 0.0157 1.0000 0.4462
-3.500 -0.4862 0.01347 0.00556 0.0177 1.0000 0.5020
-3.250 -0.4670 0.01319 0.00542 0.0194 1.0000 0.5460
-3.000 -0.4483 0.01296 0.00535 0.0214 1.0000 0.5972
-2.750 -0.4295 0.01278 0.00535 0.0235 1.0000 0.6500
-2.500 -0.4097 0.01267 0.00532 0.0253 1.0000 0.6938
-2.250 -0.3883 0.01260 0.00529 0.0267 1.0000 0.7264
-2.000 -0.3607 0.01258 0.00529 0.0268 0.9979 0.7567
-1.750 -0.3209 0.01259 0.00530 0.0245 0.9904 0.7872
-1.500 -0.2801 0.01264 0.00535 0.0220 0.9833 0.8163
-1.250 -0.2377 0.01271 0.00540 0.0192 0.9763 0.8435
-1.000 -0.1945 0.01278 0.00547 0.0163 0.9689 0.8686
-0.750 -0.1476 0.01288 0.00556 0.0126 0.9626 0.8912
-0.500 -0.1002 0.01297 0.00563 0.0088 0.9547 0.9095
-0.250 -0.0496 0.01304 0.00569 0.0043 0.9476 0.9250
0.000 0.0001 0.01306 0.00571 0.0000 0.9378 0.9378
0.250 0.0496 0.01304 0.00569 -0.0043 0.9251 0.9476
0.500 0.1002 0.01297 0.00563 -0.0088 0.9095 0.9547
0.750 0.1476 0.01288 0.00555 -0.0126 0.8911 0.9626
1.000 0.1945 0.01278 0.00547 -0.0163 0.8685 0.9689
1.250 0.2377 0.01270 0.00540 -0.0192 0.8435 0.9763
1.500 0.2800 0.01264 0.00534 -0.0220 0.8163 0.9834
1.750 0.3209 0.01259 0.00529 -0.0245 0.7875 0.9904
2.000 0.3606 0.01257 0.00528 -0.0268 0.7566 0.9979
2.250 0.3881 0.01260 0.00529 -0.0267 0.7262 1.0000
2.500 0.4096 0.01267 0.00532 -0.0253 0.6935 1.0000
2.750 0.4294 0.01278 0.00535 -0.0235 0.6501 1.0000
3.000 0.4481 0.01295 0.00535 -0.0214 0.5975 1.0000
3.250 0.4668 0.01319 0.00541 -0.0194 0.5461 1.0000
3.500 0.4860 0.01347 0.00556 -0.0176 0.5025 1.0000
3.750 0.5043 0.01383 0.00575 -0.0157 0.4473 1.0000
4.000 0.5211 0.01433 0.00594 -0.0137 0.3760 1.0000
4.250 0.5376 0.01495 0.00621 -0.0117 0.3050 1.0000
4.500 0.5559 0.01551 0.00659 -0.0101 0.2516 1.0000
4.750 0.5740 0.01616 0.00706 -0.0085 0.1965 1.0000
5.000 0.5911 0.01698 0.00760 -0.0068 0.1411 1.0000
5.250 0.6081 0.01787 0.00828 -0.0050 0.1075 1.0000
5.500 0.6252 0.01877 0.00907 -0.0033 0.0894 1.0000
5.750 0.6430 0.01959 0.00994 -0.0015 0.0784 1.0000
6.000 0.6609 0.02042 0.01086 0.0003 0.0691 1.0000
6.250 0.6781 0.02133 0.01176 0.0020 0.0616 1.0000
6.500 0.6965 0.02217 0.01273 0.0037 0.0553 1.0000
6.750 0.7132 0.02323 0.01376 0.0054 0.0495 1.0000
7.000 0.7316 0.02431 0.01501 0.0071 0.0451 1.0000
7.250 0.7491 0.02536 0.01609 0.0086 0.0408 1.0000
7.500 0.7665 0.02688 0.01771 0.0102 0.0374 1.0000
7.750 0.7851 0.02847 0.01958 0.0118 0.0348 1.0000
8.000 0.8023 0.02981 0.02106 0.0133 0.0319 1.0000
8.250 0.8176 0.03160 0.02287 0.0147 0.0296 1.0000
8.500 0.8330 0.03379 0.02542 0.0165 0.0283 1.0000
8.750 0.8455 0.03636 0.02842 0.0186 0.0271 1.0000
9.000 0.8548 0.03895 0.03143 0.0207 0.0259 1.0000
9.250 0.8609 0.04155 0.03440 0.0230 0.0247 1.0000
9.500 0.8663 0.04373 0.03684 0.0251 0.0233 1.0000
9.750 0.8683 0.04614 0.03950 0.0272 0.0226 1.0000
10.000 0.8657 0.04885 0.04246 0.0295 0.0220 1.0000
10.250 0.8573 0.05181 0.04565 0.0320 0.0218 1.0000
10.500 0.8428 0.05477 0.04882 0.0346 0.0217 1.0000
10.750 0.8243 0.05839 0.05264 0.0359 0.0217 1.0000
11.000 0.8015 0.06313 0.05761 0.0350 0.0218 1.0000
11.250 0.7696 0.07054 0.06526 0.0305 0.0223 1.0000
11.500 0.7050 0.08974 0.08465 0.0158 0.0246 1.0000
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Polar data table (+)
Polar graphs
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