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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 200,000
Max Cl/Cd: 56.2 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf25-il-200000.txt
Download as CSV file: xf-raf25-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 25 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5743   0.09440   0.09084  -0.0155   1.0000   0.0399
  -8.750  -0.5764   0.08992   0.08640  -0.0193   1.0000   0.0400
  -8.500  -0.5828   0.08474   0.08125  -0.0248   1.0000   0.0401
  -8.250  -0.5837   0.08034   0.07693  -0.0192   1.0000   0.0421
  -8.000  -0.5812   0.07734   0.07394  -0.0193   1.0000   0.0433
  -7.750  -0.5800   0.07294   0.06954  -0.0225   1.0000   0.0442
  -7.500  -0.5776   0.06840   0.06497  -0.0255   1.0000   0.0455
  -7.250  -0.5745   0.06351   0.06001  -0.0285   1.0000   0.0473
  -7.000  -0.5695   0.05837   0.05470  -0.0313   1.0000   0.0497
  -6.750  -0.5634   0.05549   0.05107  -0.0334   1.0000   0.0527
  -6.500  -0.5589   0.04793   0.04368  -0.0337   1.0000   0.0547
  -6.250  -0.5452   0.04519   0.04095  -0.0330   1.0000   0.0570
  -6.000  -0.5282   0.04524   0.04023  -0.0318   1.0000   0.0657
  -5.750  -0.4912   0.02428   0.01988  -0.0323   1.0000   0.0695
  -5.500  -0.4760   0.02176   0.01717  -0.0313   1.0000   0.0756
  -5.250  -0.4628   0.01875   0.01394  -0.0304   1.0000   0.0840
  -5.000  -0.4608   0.02551   0.01871  -0.0255   1.0000   0.0404
  -4.750  -0.4405   0.02213   0.01499  -0.0240   1.0000   0.0393
  -4.500  -0.4183   0.02019   0.01268  -0.0226   1.0000   0.0398
  -4.250  -0.3968   0.01741   0.00963  -0.0212   1.0000   0.0422
  -4.000  -0.3746   0.01638   0.00857  -0.0202   1.0000   0.0476
  -3.750  -0.3516   0.01527   0.00730  -0.0189   1.0000   0.0509
  -3.500  -0.3300   0.01400   0.00599  -0.0176   1.0000   0.0557
  -3.250  -0.3081   0.01339   0.00538  -0.0164   1.0000   0.0633
  -3.000  -0.2870   0.01255   0.00454  -0.0151   1.0000   0.0690
  -2.750  -0.2652   0.01209   0.00406  -0.0140   1.0000   0.0794
  -2.500  -0.2442   0.01142   0.00347  -0.0127   1.0000   0.0966
  -2.250  -0.2273   0.00974   0.00305  -0.0114   1.0000   0.3660
  -2.000  -0.2125   0.00890   0.00310  -0.0090   1.0000   0.5843
  -1.750  -0.1978   0.00838   0.00315  -0.0060   1.0000   0.7270
  -1.500  -0.0999   0.00789   0.00303  -0.0193   1.0000   1.0000
  -1.250  -0.0841   0.00796   0.00299  -0.0173   1.0000   1.0000
  -1.000  -0.0677   0.00806   0.00299  -0.0154   1.0000   1.0000
  -0.750  -0.0508   0.00819   0.00304  -0.0137   1.0000   1.0000
  -0.500  -0.0335   0.00835   0.00312  -0.0121   1.0000   1.0000
  -0.250  -0.0016   0.00852   0.00322  -0.0135   0.9968   1.0000
   0.000   0.0430   0.00866   0.00331  -0.0174   0.9898   1.0000
   0.250   0.0906   0.00878   0.00341  -0.0218   0.9831   1.0000
   0.500   0.1368   0.00878   0.00341  -0.0258   0.9731   1.0000
   0.750   0.1827   0.00874   0.00340  -0.0296   0.9637   1.0000
   1.000   0.2307   0.00868   0.00339  -0.0338   0.9569   1.0000
   1.250   0.2770   0.00852   0.00331  -0.0375   0.9456   1.0000
   1.500   0.3289   0.00812   0.00298  -0.0416   0.9217   1.0000
   1.750   0.3701   0.00791   0.00280  -0.0435   0.8921   1.0000
   2.000   0.4007   0.00788   0.00277  -0.0434   0.8610   1.0000
   2.250   0.4249   0.00794   0.00278  -0.0417   0.8183   1.0000
   2.500   0.4465   0.00808   0.00279  -0.0396   0.7641   1.0000
   2.750   0.4670   0.00831   0.00283  -0.0374   0.6938   1.0000
   3.000   0.4837   0.00885   0.00288  -0.0344   0.5714   1.0000
   3.250   0.4899   0.01063   0.00321  -0.0304   0.2663   1.0000
   3.500   0.5043   0.01237   0.00402  -0.0283   0.0892   1.0000
   3.750   0.5260   0.01314   0.00472  -0.0271   0.0699   1.0000
   4.000   0.5457   0.01428   0.00584  -0.0255   0.0593   1.0000
   4.250   0.5670   0.01529   0.00685  -0.0242   0.0494   1.0000
   4.500   0.5883   0.01687   0.00846  -0.0227   0.0449   1.0000
   4.750   0.6125   0.01813   0.00986  -0.0216   0.0413   1.0000
   5.000   0.6358   0.01943   0.01114  -0.0208   0.0361   1.0000
   5.250   0.6596   0.02303   0.01495  -0.0198   0.0343   1.0000
   5.500   0.6833   0.02506   0.01728  -0.0186   0.0338   1.0000
   5.750   0.7068   0.02776   0.02026  -0.0173   0.0347   1.0000
   6.000   0.7278   0.03052   0.02337  -0.0159   0.0348   1.0000
   6.250   0.7471   0.03360   0.02681  -0.0142   0.0348   1.0000
   7.250   0.8054   0.04918   0.04441  -0.0054   0.0495   1.0000
   7.500   0.8182   0.05387   0.04903  -0.0051   0.0480   1.0000
   7.750   0.8126   0.06240   0.05778  -0.0046   0.0468   1.0000
   8.000   0.8140   0.06478   0.06055  -0.0030   0.0464   1.0000
   8.250   0.8072   0.06688   0.06316  -0.0018   0.0441   1.0000
   8.500   0.7986   0.07144   0.06790  -0.0021   0.0429   1.0000
   8.750   0.7859   0.07581   0.07236  -0.0025   0.0425   1.0000
   9.000   0.7694   0.08074   0.07735  -0.0047   0.0424   1.0000
   9.250   0.7553   0.08759   0.08419  -0.0110   0.0426   1.0000
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