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RAF 27 AIRFOIL (raf27-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAF 27 AIRFOIL (raf27-il)
Reynolds number: 50,000
Max Cl/Cd: 28.27 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf27-il-50000.txt
Download as CSV file: xf-raf27-il-50000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5863   0.08337   0.07670  -0.0184   1.0000   0.1653
  -9.250  -0.5968   0.07513   0.06850  -0.0214   1.0000   0.1533
  -9.000  -0.6169   0.06802   0.06140  -0.0242   1.0000   0.1460
  -8.750  -0.7443   0.06948   0.06240  -0.0227   1.0000   0.1338
  -8.500  -0.7586   0.06393   0.05665  -0.0223   1.0000   0.1289
  -8.250  -0.7838   0.05822   0.05027  -0.0207   1.0000   0.1227
  -8.000  -0.7790   0.05377   0.04557  -0.0192   1.0000   0.1212
  -7.750  -0.7759   0.04974   0.04111  -0.0172   1.0000   0.1211
  -7.500  -0.7716   0.04610   0.03689  -0.0148   1.0000   0.1225
  -7.250  -0.7609   0.04255   0.03291  -0.0128   1.0000   0.1249
  -7.000  -0.7433   0.03940   0.02957  -0.0113   1.0000   0.1280
  -6.750  -0.7261   0.03664   0.02643  -0.0096   1.0000   0.1329
  -6.500  -0.7081   0.03409   0.02354  -0.0078   1.0000   0.1406
  -6.250  -0.6861   0.03176   0.02092  -0.0064   1.0000   0.1493
  -6.000  -0.6637   0.02970   0.01878  -0.0051   1.0000   0.1630
  -5.750  -0.6397   0.02770   0.01671  -0.0039   1.0000   0.1806
  -5.500  -0.6166   0.02581   0.01498  -0.0025   1.0000   0.2092
  -5.250  -0.5975   0.02384   0.01347  -0.0005   1.0000   0.2620
  -5.000  -0.5881   0.02165   0.01227   0.0034   1.0000   0.3750
  -4.750  -0.5831   0.02063   0.01201   0.0089   1.0000   0.5001
  -4.500  -0.5733   0.02046   0.01228   0.0145   1.0000   0.5900
  -4.250  -0.5593   0.02060   0.01255   0.0198   1.0000   0.6596
  -4.000  -0.5399   0.02081   0.01280   0.0243   1.0000   0.7173
  -3.750  -0.5128   0.02146   0.01340   0.0284   1.0000   0.7735
  -3.500  -0.4620   0.02264   0.01428   0.0289   1.0000   0.8287
  -3.250  -0.3847   0.02343   0.01457   0.0225   1.0000   0.8718
  -3.000  -0.2891   0.02370   0.01426   0.0111   1.0000   0.9093
  -2.750  -0.2074   0.02328   0.01345   0.0007   1.0000   0.9434
  -2.500  -0.1226   0.02231   0.01212  -0.0113   1.0000   0.9750
  -2.250  -0.0501   0.02103   0.01059  -0.0221   1.0000   1.0000
  -2.000  -0.0406   0.02039   0.00992  -0.0209   1.0000   1.0000
  -1.750  -0.0316   0.01983   0.00935  -0.0194   1.0000   1.0000
  -1.500  -0.0233   0.01935   0.00887  -0.0176   1.0000   1.0000
  -1.250  -0.0158   0.01894   0.00846  -0.0155   1.0000   1.0000
  -1.000  -0.0095   0.01860   0.00814  -0.0131   1.0000   1.0000
  -0.750  -0.0044   0.01833   0.00789  -0.0104   1.0000   1.0000
  -0.500  -0.0011   0.01812   0.00771  -0.0073   1.0000   1.0000
  -0.250   0.0002   0.01800   0.00761  -0.0038   1.0000   1.0000
   0.000   0.0000   0.01795   0.00757   0.0000   1.0000   1.0000
   0.250  -0.0002   0.01800   0.00761   0.0038   1.0000   1.0000
   0.500   0.0011   0.01812   0.00771   0.0073   1.0000   1.0000
   0.750   0.0044   0.01832   0.00789   0.0104   1.0000   1.0000
   1.000   0.0095   0.01860   0.00814   0.0131   1.0000   1.0000
   1.250   0.0158   0.01894   0.00846   0.0155   1.0000   1.0000
   1.500   0.0233   0.01935   0.00886   0.0176   1.0000   1.0000
   1.750   0.0316   0.01983   0.00935   0.0194   1.0000   1.0000
   2.000   0.0406   0.02038   0.00992   0.0209   1.0000   1.0000
   2.250   0.0501   0.02102   0.01059   0.0221   1.0000   1.0000
   2.500   0.1224   0.02230   0.01211   0.0114   0.9751   1.0000
   2.750   0.2075   0.02327   0.01344  -0.0007   0.9434   1.0000
   3.000   0.2892   0.02370   0.01426  -0.0111   0.9092   1.0000
   3.250   0.3845   0.02344   0.01457  -0.0225   0.8718   1.0000
   3.500   0.4623   0.02263   0.01427  -0.0289   0.8286   1.0000
   3.750   0.5129   0.02146   0.01340  -0.0284   0.7734   1.0000
   4.000   0.5399   0.02081   0.01280  -0.0243   0.7173   1.0000
   4.250   0.5594   0.02060   0.01255  -0.0198   0.6596   1.0000
   4.500   0.5734   0.02047   0.01229  -0.0146   0.5900   1.0000
   4.750   0.5832   0.02063   0.01201  -0.0089   0.5002   1.0000
   5.000   0.5880   0.02165   0.01227  -0.0034   0.3747   1.0000
   5.250   0.5976   0.02384   0.01347   0.0005   0.2621   1.0000
   5.500   0.6167   0.02582   0.01499   0.0025   0.2090   1.0000
   5.750   0.6398   0.02770   0.01671   0.0039   0.1806   1.0000
   6.000   0.6636   0.02969   0.01878   0.0051   0.1628   1.0000
   6.250   0.6861   0.03176   0.02093   0.0064   0.1492   1.0000
   6.500   0.7082   0.03410   0.02356   0.0078   0.1407   1.0000
   6.750   0.7261   0.03665   0.02645   0.0096   0.1328   1.0000
   7.000   0.7433   0.03941   0.02957   0.0113   0.1279   1.0000
   7.250   0.7611   0.04256   0.03289   0.0128   0.1248   1.0000
   7.500   0.7714   0.04608   0.03687   0.0149   0.1224   1.0000
   7.750   0.7757   0.04973   0.04110   0.0173   0.1210   1.0000
   8.000   0.7788   0.05378   0.04558   0.0192   0.1212   1.0000
   8.250   0.7836   0.05822   0.05027   0.0207   0.1227   1.0000
   8.500   0.7585   0.06393   0.05666   0.0223   0.1289   1.0000
   8.750   0.7443   0.06947   0.06239   0.0227   0.1337   1.0000
   9.000   0.7242   0.07530   0.06836   0.0220   0.1403   1.0000
   9.250   0.6869   0.08213   0.07515   0.0196   0.1481   1.0000
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