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AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-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-rg12a-il-50000.txt
Download as CSV file: xf-rg12a-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AIRFOIL PROFILE12A 9.00%                        
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4618   0.10713   0.10013  -0.0110   1.0000   0.2412
  -8.750  -0.4612   0.10397   0.09705  -0.0110   1.0000   0.2516
  -8.500  -0.4838   0.10315   0.09639  -0.0124   1.0000   0.2611
  -8.250  -0.4567   0.09777   0.09097  -0.0098   1.0000   0.2800
  -8.000  -0.4518   0.09466   0.08793  -0.0088   1.0000   0.2966
  -7.750  -0.4500   0.09183   0.08517  -0.0077   1.0000   0.3132
  -7.500  -0.4635   0.09008   0.08356  -0.0065   1.0000   0.3317
  -7.250  -0.4519   0.08624   0.07976  -0.0048   1.0000   0.3494
  -7.000  -0.4423   0.08290   0.07643  -0.0030   1.0000   0.3678
  -6.750  -0.4665   0.08210   0.07582   0.0000   1.0000   0.3896
  -5.750  -0.4915   0.05073   0.04331  -0.0352   1.0000   0.1312
  -5.500  -0.4766   0.04575   0.03765  -0.0355   1.0000   0.1187
  -5.250  -0.4594   0.04189   0.03355  -0.0347   1.0000   0.1150
  -5.000  -0.4397   0.03787   0.02888  -0.0341   1.0000   0.1094
  -4.750  -0.4169   0.03464   0.02480  -0.0331   1.0000   0.1067
  -4.500  -0.3949   0.03184   0.02186  -0.0321   1.0000   0.1099
  -4.250  -0.3714   0.02973   0.01927  -0.0310   1.0000   0.1183
  -4.000  -0.3474   0.02742   0.01682  -0.0298   1.0000   0.1272
  -3.750  -0.3231   0.02551   0.01474  -0.0285   1.0000   0.1445
  -3.500  -0.2989   0.02361   0.01287  -0.0270   1.0000   0.1769
  -3.250  -0.2769   0.02158   0.01135  -0.0254   1.0000   0.2506
  -3.000  -0.1105   0.01900   0.01072  -0.0364   1.0000   1.0000
  -2.750  -0.1061   0.01861   0.01013  -0.0337   1.0000   1.0000
  -2.500  -0.1020   0.01831   0.00965  -0.0308   1.0000   1.0000
  -2.250  -0.0985   0.01806   0.00925  -0.0276   1.0000   1.0000
  -2.000  -0.0957   0.01786   0.00887  -0.0243   1.0000   1.0000
  -1.750  -0.0932   0.01770   0.00858  -0.0208   1.0000   1.0000
  -1.500  -0.0899   0.01759   0.00833  -0.0176   1.0000   1.0000
  -1.250  -0.0836   0.01754   0.00813  -0.0148   1.0000   1.0000
  -1.000  -0.0729   0.01757   0.00799  -0.0129   1.0000   1.0000
  -0.750  -0.0591   0.01769   0.00793  -0.0115   1.0000   1.0000
  -0.500  -0.0433   0.01787   0.00792  -0.0105   1.0000   1.0000
  -0.250  -0.0265   0.01811   0.00802  -0.0097   1.0000   1.0000
   0.000  -0.0092   0.01841   0.00819  -0.0091   1.0000   1.0000
   0.250   0.0084   0.01878   0.00844  -0.0087   1.0000   1.0000
   0.500   0.0259   0.01921   0.00877  -0.0083   1.0000   1.0000
   0.750   0.0433   0.01970   0.00916  -0.0080   1.0000   1.0000
   1.000   0.0606   0.02026   0.00965  -0.0078   1.0000   1.0000
   1.250   0.0776   0.02088   0.01022  -0.0078   1.0000   1.0000
   1.500   0.0942   0.02157   0.01087  -0.0078   1.0000   1.0000
   1.750   0.1318   0.02267   0.01196  -0.0118   0.9899   1.0000
   2.000   0.1922   0.02402   0.01333  -0.0196   0.9687   1.0000
   2.250   0.2468   0.02501   0.01438  -0.0260   0.9462   1.0000
   2.500   0.2986   0.02579   0.01526  -0.0315   0.9231   1.0000
   2.750   0.3565   0.02640   0.01601  -0.0374   0.9009   1.0000
   3.000   0.4045   0.02678   0.01654  -0.0413   0.8764   1.0000
   3.250   0.4531   0.02700   0.01697  -0.0448   0.8519   1.0000
   3.500   0.5117   0.02680   0.01703  -0.0492   0.8282   1.0000
   3.750   0.5699   0.02628   0.01683  -0.0528   0.8039   1.0000
   4.000   0.6154   0.02586   0.01664  -0.0540   0.7769   1.0000
   4.250   0.6594   0.02531   0.01631  -0.0544   0.7485   1.0000
   4.500   0.7014   0.02469   0.01593  -0.0541   0.7182   1.0000
   4.750   0.7298   0.02462   0.01600  -0.0523   0.6835   1.0000
   5.000   0.7630   0.02429   0.01575  -0.0506   0.6477   1.0000
   5.250   0.7876   0.02439   0.01591  -0.0482   0.6081   1.0000
   5.500   0.8118   0.02453   0.01609  -0.0456   0.5662   1.0000
   5.750   0.8349   0.02475   0.01625  -0.0430   0.5215   1.0000
   6.000   0.8550   0.02516   0.01657  -0.0402   0.4734   1.0000
   6.250   0.8740   0.02571   0.01693  -0.0374   0.4216   1.0000
   6.500   0.8899   0.02654   0.01755  -0.0344   0.3665   1.0000
   6.750   0.9046   0.02769   0.01846  -0.0314   0.3101   1.0000
   7.000   0.9177   0.02918   0.01964  -0.0285   0.2546   1.0000
   7.250   0.9298   0.03095   0.02113  -0.0258   0.2044   1.0000
   7.500   0.9449   0.03319   0.02313  -0.0237   0.1645   1.0000
   7.750   0.9633   0.03571   0.02558  -0.0222   0.1378   1.0000
   8.000   0.9851   0.03888   0.02897  -0.0209   0.1232   1.0000
   8.250   1.0071   0.04224   0.03241  -0.0201   0.1135   1.0000
   8.500   1.0190   0.04535   0.03613  -0.0179   0.1065   1.0000
   8.750   1.0373   0.04906   0.03993  -0.0171   0.1008   1.0000
   9.000   1.0423   0.05320   0.04472  -0.0148   0.0997   1.0000
   9.250   1.0434   0.05766   0.04971  -0.0126   0.0993   1.0000
   9.500   1.0415   0.06230   0.05478  -0.0107   0.0996   1.0000
   9.750   1.0365   0.06711   0.05992  -0.0091   0.1002   1.0000
  10.000   1.0300   0.07209   0.06515  -0.0078   0.1008   1.0000
  10.250   0.9840   0.07641   0.06994  -0.0055   0.1038   1.0000
  10.500   0.9356   0.08268   0.07638  -0.0069   0.1071   1.0000
  10.750   0.9028   0.09032   0.08401  -0.0113   0.1101   1.0000
  11.000   0.8861   0.09783   0.09150  -0.0153   0.1125   1.0000
  11.250   0.8877   0.10381   0.09748  -0.0166   0.1142   1.0000
  11.500   0.8100   0.12623   0.11959  -0.0376   0.1489   1.0000
  11.750   0.8245   0.13201   0.12545  -0.0376   0.1552   1.0000
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