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RAF 30 MOD AIRFOIL (raf30md-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: RAF 30 MOD AIRFOIL (raf30md-il)
Reynolds number: 200,000
Max Cl/Cd: 36.31 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf30md-il-200000-n5.txt
Download as CSV file: xf-raf30md-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 30 MOD 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6978   0.08623   0.08279  -0.0012   1.0000   0.0103
  -9.750  -0.7257   0.07102   0.06765  -0.0125   1.0000   0.0097
  -9.500  -0.7610   0.05940   0.05584  -0.0202   1.0000   0.0092
  -9.250  -0.7824   0.05452   0.05077  -0.0201   1.0000   0.0091
  -9.000  -0.7953   0.04963   0.04561  -0.0194   1.0000   0.0091
  -8.750  -0.8027   0.04496   0.04060  -0.0183   1.0000   0.0091
  -8.500  -0.8037   0.04066   0.03592  -0.0169   1.0000   0.0093
  -8.250  -0.7930   0.03874   0.03381  -0.0161   1.0000   0.0098
  -8.000  -0.7815   0.03632   0.03111  -0.0149   1.0000   0.0103
  -7.750  -0.7666   0.03441   0.02894  -0.0139   1.0000   0.0112
  -7.500  -0.7538   0.03120   0.02530  -0.0123   1.0000   0.0117
  -7.250  -0.7386   0.02818   0.02182  -0.0107   1.0000   0.0122
  -7.000  -0.7206   0.02566   0.01890  -0.0093   1.0000   0.0128
  -6.750  -0.7007   0.02354   0.01643  -0.0080   1.0000   0.0135
  -6.500  -0.6786   0.02227   0.01489  -0.0069   1.0000   0.0144
  -6.250  -0.6596   0.02000   0.01242  -0.0057   1.0000   0.0164
  -6.000  -0.6377   0.01890   0.01122  -0.0047   1.0000   0.0182
  -5.750  -0.6158   0.01778   0.00993  -0.0036   1.0000   0.0204
  -5.500  -0.5931   0.01699   0.00900  -0.0025   1.0000   0.0239
  -5.250  -0.5701   0.01633   0.00823  -0.0016   1.0000   0.0264
  -5.000  -0.5503   0.01518   0.00706  -0.0003   1.0000   0.0307
  -4.750  -0.5276   0.01465   0.00644   0.0006   1.0000   0.0357
  -4.500  -0.5061   0.01393   0.00562   0.0018   1.0000   0.0398
  -4.250  -0.4835   0.01343   0.00507   0.0027   1.0000   0.0451
  -4.000  -0.4605   0.01302   0.00457   0.0036   1.0000   0.0490
  -3.750  -0.4380   0.01251   0.00405   0.0046   1.0000   0.0565
  -3.500  -0.4154   0.01204   0.00361   0.0056   1.0000   0.0721
  -3.250  -0.3939   0.01142   0.00323   0.0066   1.0000   0.1276
  -3.000  -0.3732   0.01074   0.00290   0.0076   1.0000   0.2110
  -2.750  -0.3529   0.01009   0.00266   0.0087   1.0000   0.3103
  -2.500  -0.3331   0.00950   0.00249   0.0100   1.0000   0.4120
  -2.250  -0.3127   0.00910   0.00237   0.0113   1.0000   0.4900
  -2.000  -0.2922   0.00880   0.00232   0.0127   1.0000   0.5574
  -1.750  -0.2602   0.00847   0.00233   0.0118   0.9930   0.6412
  -1.500  -0.2285   0.00816   0.00238   0.0112   0.9849   0.7318
  -1.250  -0.1956   0.00798   0.00243   0.0106   0.9765   0.7973
  -1.000  -0.1597   0.00789   0.00243   0.0092   0.9681   0.8314
  -0.750  -0.1224   0.00785   0.00241   0.0075   0.9581   0.8567
  -0.500  -0.0837   0.00782   0.00240   0.0054   0.9469   0.8805
  -0.250  -0.0423   0.00781   0.00240   0.0028   0.9343   0.9008
   0.000   0.0000   0.00781   0.00241   0.0000   0.9189   0.9189
   0.250   0.0424   0.00781   0.00241  -0.0028   0.9008   0.9344
   0.500   0.0838   0.00782   0.00240  -0.0054   0.8803   0.9468
   0.750   0.1224   0.00785   0.00241  -0.0075   0.8567   0.9581
   1.000   0.1597   0.00790   0.00243  -0.0092   0.8314   0.9681
   1.250   0.1956   0.00798   0.00243  -0.0106   0.7972   0.9765
   1.500   0.2285   0.00816   0.00239  -0.0112   0.7323   0.9848
   1.750   0.2604   0.00846   0.00233  -0.0118   0.6451   0.9929
   2.000   0.2924   0.00879   0.00233  -0.0128   0.5588   1.0000
   2.250   0.3128   0.00910   0.00237  -0.0113   0.4883   1.0000
   2.750   0.3531   0.01009   0.00266  -0.0087   0.3111   1.0000
   3.000   0.3733   0.01075   0.00291  -0.0076   0.2085   1.0000
   3.250   0.3941   0.01142   0.00323  -0.0066   0.1285   1.0000
   3.500   0.4155   0.01204   0.00363  -0.0056   0.0719   1.0000
   3.750   0.4381   0.01251   0.00405  -0.0046   0.0565   1.0000
   4.000   0.4606   0.01302   0.00457  -0.0036   0.0490   1.0000
   4.250   0.4837   0.01343   0.00507  -0.0027   0.0451   1.0000
   4.500   0.5062   0.01394   0.00563  -0.0018   0.0398   1.0000
   4.750   0.5277   0.01465   0.00644  -0.0007   0.0357   1.0000
   5.000   0.5504   0.01519   0.00707   0.0003   0.0307   1.0000
   5.250   0.5700   0.01637   0.00826   0.0016   0.0263   1.0000
   5.500   0.5931   0.01700   0.00902   0.0025   0.0240   1.0000
   5.750   0.6158   0.01778   0.00993   0.0036   0.0203   1.0000
   6.000   0.6377   0.01890   0.01122   0.0047   0.0183   1.0000
   6.250   0.6596   0.02001   0.01244   0.0057   0.0165   1.0000
   6.500   0.6784   0.02229   0.01491   0.0069   0.0144   1.0000
   6.750   0.7006   0.02357   0.01645   0.0080   0.0135   1.0000
   7.000   0.7205   0.02565   0.01888   0.0093   0.0128   1.0000
   7.250   0.7385   0.02817   0.02181   0.0107   0.0122   1.0000
   7.500   0.7536   0.03122   0.02532   0.0124   0.0117   1.0000
   7.750   0.7662   0.03447   0.02901   0.0140   0.0113   1.0000
   8.000   0.7797   0.03678   0.03162   0.0151   0.0105   1.0000
   8.250   0.7910   0.03925   0.03437   0.0163   0.0099   1.0000
   8.500   0.8020   0.04109   0.03640   0.0171   0.0094   1.0000
   8.750   0.8026   0.04490   0.04054   0.0184   0.0091   1.0000
   9.000   0.7955   0.04960   0.04558   0.0195   0.0091   1.0000
   9.250   0.7821   0.05443   0.05068   0.0201   0.0091   1.0000
   9.500   0.7595   0.05980   0.05624   0.0202   0.0093   1.0000
   9.750   0.7270   0.07047   0.06709   0.0129   0.0096   1.0000
  10.000   0.6947   0.08721   0.08377   0.0006   0.0107   1.0000
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