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RG 12A-1.8/9.0 AIRFOIL (rg12a189-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RG 12A-1.8/9.0 AIRFOIL (rg12a189-il)
Reynolds number: 50,000
Max Cl/Cd: 33.99 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rg12a189-il-50000.txt
Download as CSV file: xf-rg12a189-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 12A-1.8/9.0 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4615   0.10399   0.09706  -0.0110   1.0000   0.2515
  -8.500  -0.4651   0.10124   0.09441  -0.0111   1.0000   0.2636
  -8.250  -0.4797   0.09980   0.09309  -0.0114   1.0000   0.2753
  -8.000  -0.4517   0.09469   0.08795  -0.0087   1.0000   0.2970
  -7.750  -0.4514   0.09193   0.08527  -0.0076   1.0000   0.3142
  -7.500  -0.4624   0.09000   0.08347  -0.0065   1.0000   0.3316
  -7.250  -0.4518   0.08625   0.07976  -0.0048   1.0000   0.3495
  -7.000  -0.4424   0.08289   0.07641  -0.0030   1.0000   0.3682
  -6.750  -0.4675   0.08214   0.07585   0.0001   1.0000   0.3896
  -6.500  -0.4400   0.07732   0.07101   0.0014   1.0000   0.4087
  -5.750  -0.4914   0.05069   0.04327  -0.0352   1.0000   0.1313
  -5.500  -0.4764   0.04574   0.03763  -0.0355   1.0000   0.1186
  -5.250  -0.4592   0.04187   0.03352  -0.0347   1.0000   0.1150
  -5.000  -0.4394   0.03785   0.02885  -0.0341   1.0000   0.1094
  -4.750  -0.4166   0.03461   0.02478  -0.0331   1.0000   0.1068
  -4.500  -0.3946   0.03184   0.02186  -0.0321   1.0000   0.1100
  -4.250  -0.3710   0.02972   0.01924  -0.0310   1.0000   0.1185
  -4.000  -0.3470   0.02741   0.01681  -0.0299   1.0000   0.1272
  -3.750  -0.3227   0.02550   0.01473  -0.0286   1.0000   0.1447
  -3.500  -0.2985   0.02361   0.01287  -0.0270   1.0000   0.1769
  -3.250  -0.2765   0.02157   0.01134  -0.0255   1.0000   0.2509
  -3.000  -0.1106   0.01899   0.01072  -0.0364   1.0000   1.0000
  -2.750  -0.1062   0.01861   0.01013  -0.0337   1.0000   1.0000
  -2.500  -0.1020   0.01830   0.00964  -0.0308   1.0000   1.0000
  -2.250  -0.0985   0.01806   0.00924  -0.0276   1.0000   1.0000
  -2.000  -0.0957   0.01785   0.00887  -0.0243   1.0000   1.0000
  -1.750  -0.0931   0.01769   0.00857  -0.0209   1.0000   1.0000
  -1.500  -0.0898   0.01758   0.00832  -0.0176   1.0000   1.0000
  -1.250  -0.0835   0.01753   0.00812  -0.0148   1.0000   1.0000
  -1.000  -0.0727   0.01757   0.00798  -0.0129   1.0000   1.0000
  -0.750  -0.0588   0.01768   0.00792  -0.0115   1.0000   1.0000
  -0.500  -0.0430   0.01786   0.00792  -0.0105   1.0000   1.0000
  -0.250  -0.0262   0.01810   0.00801  -0.0098   1.0000   1.0000
   0.000  -0.0088   0.01841   0.00819  -0.0092   1.0000   1.0000
   0.250   0.0087   0.01878   0.00844  -0.0087   1.0000   1.0000
   0.500   0.0263   0.01920   0.00877  -0.0083   1.0000   1.0000
   0.750   0.0437   0.01969   0.00916  -0.0081   1.0000   1.0000
   1.000   0.0609   0.02025   0.00965  -0.0079   1.0000   1.0000
   1.250   0.0779   0.02088   0.01022  -0.0078   1.0000   1.0000
   1.500   0.0945   0.02157   0.01087  -0.0078   1.0000   1.0000
   1.750   0.1323   0.02268   0.01196  -0.0119   0.9898   1.0000
   2.000   0.1925   0.02402   0.01333  -0.0197   0.9688   1.0000
   2.250   0.2473   0.02500   0.01437  -0.0261   0.9462   1.0000
   2.500   0.2991   0.02578   0.01526  -0.0315   0.9230   1.0000
   2.750   0.3568   0.02639   0.01600  -0.0375   0.9009   1.0000
   3.000   0.4050   0.02677   0.01653  -0.0414   0.8764   1.0000
   3.250   0.4535   0.02699   0.01697  -0.0449   0.8518   1.0000
   3.500   0.5120   0.02680   0.01702  -0.0493   0.8282   1.0000
   3.750   0.5703   0.02627   0.01681  -0.0528   0.8039   1.0000
   4.000   0.6158   0.02585   0.01663  -0.0540   0.7769   1.0000
   4.250   0.6598   0.02529   0.01630  -0.0545   0.7484   1.0000
   4.500   0.7018   0.02467   0.01592  -0.0542   0.7181   1.0000
   4.750   0.7300   0.02461   0.01599  -0.0523   0.6834   1.0000
   5.000   0.7633   0.02428   0.01574  -0.0506   0.6476   1.0000
   5.250   0.7878   0.02439   0.01591  -0.0482   0.6081   1.0000
   5.500   0.8120   0.02453   0.01609  -0.0457   0.5662   1.0000
   5.750   0.8350   0.02475   0.01624  -0.0430   0.5214   1.0000
   6.000   0.8550   0.02517   0.01658  -0.0402   0.4732   1.0000
   6.250   0.8741   0.02572   0.01694  -0.0374   0.4216   1.0000
   6.500   0.8900   0.02654   0.01755  -0.0344   0.3665   1.0000
   6.750   0.9047   0.02769   0.01846  -0.0314   0.3100   1.0000
   7.000   0.9178   0.02918   0.01964  -0.0286   0.2545   1.0000
   7.250   0.9298   0.03095   0.02114  -0.0258   0.2042   1.0000
   7.500   0.9450   0.03319   0.02313  -0.0237   0.1646   1.0000
   7.750   0.9634   0.03571   0.02558  -0.0222   0.1377   1.0000
   8.000   0.9852   0.03889   0.02898  -0.0209   0.1231   1.0000
   8.250   1.0072   0.04225   0.03241  -0.0201   0.1135   1.0000
   8.500   1.0191   0.04534   0.03613  -0.0179   0.1064   1.0000
   8.750   1.0373   0.04908   0.03995  -0.0171   0.1007   1.0000
   9.000   1.0425   0.05320   0.04471  -0.0148   0.0997   1.0000
   9.250   1.0436   0.05767   0.04971  -0.0127   0.0994   1.0000
   9.500   1.0416   0.06231   0.05478  -0.0107   0.0996   1.0000
   9.750   1.0365   0.06713   0.05994  -0.0091   0.1001   1.0000
  10.000   1.0300   0.07208   0.06514  -0.0078   0.1007   1.0000
  10.250   0.9844   0.07641   0.06994  -0.0056   0.1038   1.0000
  10.500   0.9357   0.08269   0.07639  -0.0069   0.1071   1.0000
  10.750   0.9029   0.09032   0.08400  -0.0113   0.1101   1.0000
  11.000   0.8856   0.09795   0.09161  -0.0155   0.1125   1.0000
  11.250   0.8884   0.10376   0.09743  -0.0165   0.1142   1.0000
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