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PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il)
Reynolds number: 200,000
Max Cl/Cd: 56.31 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-pfcm-il-200000-n5.txt
Download as CSV file: xf-pfcm-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: PROPFAN CRUISE MISSILE WING 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5764   0.08687   0.08343  -0.0156   1.0000   0.0129
  -8.750  -0.5782   0.08146   0.07807  -0.0196   1.0000   0.0129
  -8.500  -0.5803   0.07488   0.07152  -0.0265   1.0000   0.0128
  -8.250  -0.5818   0.06851   0.06508  -0.0325   1.0000   0.0128
  -8.000  -0.5834   0.06281   0.05927  -0.0360   1.0000   0.0129
  -7.750  -0.5818   0.05797   0.05428  -0.0382   1.0000   0.0132
  -7.500  -0.5742   0.05504   0.05127  -0.0391   1.0000   0.0139
  -7.250  -0.5694   0.04714   0.04295  -0.0403   1.0000   0.0130
  -7.000  -0.5592   0.04227   0.03774  -0.0401   1.0000   0.0128
  -6.750  -0.5483   0.03778   0.03288  -0.0394   1.0000   0.0128
  -6.500  -0.5364   0.03359   0.02828  -0.0385   1.0000   0.0130
  -6.250  -0.5220   0.03023   0.02455  -0.0374   1.0000   0.0133
  -6.000  -0.5051   0.02768   0.02167  -0.0363   1.0000   0.0138
  -5.750  -0.4867   0.02581   0.01952  -0.0351   1.0000   0.0146
  -5.500  -0.4672   0.02467   0.01820  -0.0341   1.0000   0.0164
  -5.250  -0.4472   0.02277   0.01594  -0.0328   1.0000   0.0176
  -5.000  -0.4263   0.02084   0.01366  -0.0315   1.0000   0.0182
  -4.750  -0.3934   0.01904   0.01150  -0.0325   0.9969   0.0191
  -4.500  -0.3595   0.01798   0.01021  -0.0339   0.9934   0.0202
  -4.250  -0.3279   0.01619   0.00829  -0.0351   0.9898   0.0223
  -4.000  -0.2942   0.01507   0.00708  -0.0365   0.9865   0.0234
  -3.750  -0.2616   0.01419   0.00614  -0.0378   0.9817   0.0249
  -3.500  -0.2273   0.01346   0.00530  -0.0394   0.9777   0.0273
  -3.250  -0.1945   0.01289   0.00461  -0.0407   0.9723   0.0305
  -3.000  -0.1611   0.01232   0.00402  -0.0421   0.9670   0.0398
  -2.750  -0.1289   0.01179   0.00353  -0.0432   0.9610   0.0627
  -2.500  -0.0985   0.01063   0.00308  -0.0447   0.9547   0.2389
  -2.250  -0.0724   0.00931   0.00297  -0.0451   0.9478   0.5539
  -2.000  -0.0434   0.00903   0.00294  -0.0452   0.9406   0.6424
  -1.750  -0.0175   0.00881   0.00299  -0.0444   0.9322   0.7295
  -1.500   0.0094   0.00864   0.00294  -0.0437   0.9249   0.7832
  -1.250   0.0363   0.00853   0.00284  -0.0432   0.9153   0.8033
  -1.000   0.0640   0.00844   0.00274  -0.0430   0.9064   0.8200
  -0.750   0.0917   0.00836   0.00264  -0.0427   0.8979   0.8373
  -0.500   0.1182   0.00829   0.00258  -0.0422   0.8878   0.8572
  -0.250   0.1458   0.00823   0.00253  -0.0418   0.8783   0.8809
   0.250   0.2094   0.00814   0.00246  -0.0431   0.8596   0.9407
   0.500   0.2460   0.00812   0.00243  -0.0450   0.8501   0.9747
   0.750   0.2802   0.00812   0.00240  -0.0464   0.8402   1.0000
   1.000   0.3067   0.00816   0.00241  -0.0462   0.8290   1.0000
   1.250   0.3334   0.00821   0.00246  -0.0460   0.8173   1.0000
   1.500   0.3602   0.00828   0.00251  -0.0458   0.8059   1.0000
   1.750   0.3870   0.00835   0.00258  -0.0456   0.7944   1.0000
   2.000   0.4138   0.00842   0.00266  -0.0453   0.7820   1.0000
   2.250   0.4402   0.00851   0.00277  -0.0449   0.7655   1.0000
   2.500   0.4652   0.00861   0.00278  -0.0440   0.7349   1.0000
   2.750   0.4884   0.00879   0.00274  -0.0426   0.6774   1.0000
   3.000   0.5119   0.00909   0.00279  -0.0414   0.6122   1.0000
   3.250   0.5349   0.00953   0.00294  -0.0404   0.5315   1.0000
   3.500   0.5500   0.01108   0.00332  -0.0386   0.2867   1.0000
   3.750   0.5682   0.01265   0.00400  -0.0378   0.1035   1.0000
   4.000   0.5917   0.01343   0.00454  -0.0373   0.0547   1.0000
   4.250   0.6160   0.01411   0.00517  -0.0368   0.0390   1.0000
   4.500   0.6408   0.01468   0.00588  -0.0364   0.0337   1.0000
   4.750   0.6646   0.01540   0.00665  -0.0358   0.0292   1.0000
   5.000   0.6880   0.01619   0.00751  -0.0352   0.0263   1.0000
   5.250   0.7113   0.01701   0.00844  -0.0345   0.0247   1.0000
   5.500   0.7342   0.01798   0.00951  -0.0337   0.0232   1.0000
   5.750   0.7570   0.01909   0.01070  -0.0328   0.0219   1.0000
   6.000   0.7804   0.02008   0.01180  -0.0322   0.0200   1.0000
   6.250   0.8018   0.02190   0.01371  -0.0314   0.0178   1.0000
   6.500   0.8254   0.02360   0.01560  -0.0306   0.0170   1.0000
   6.750   0.8491   0.02533   0.01760  -0.0298   0.0163   1.0000
   7.000   0.8720   0.02715   0.01973  -0.0289   0.0151   1.0000
   7.250   0.8941   0.02857   0.02140  -0.0282   0.0134   1.0000
   7.500   0.9142   0.03061   0.02375  -0.0272   0.0125   1.0000
   7.750   0.9321   0.03293   0.02645  -0.0260   0.0119   1.0000
   8.000   0.9473   0.03559   0.02944  -0.0247   0.0114   1.0000
   8.250   0.9576   0.03916   0.03340  -0.0230   0.0110   1.0000
   8.500   0.9612   0.04377   0.03848  -0.0208   0.0108   1.0000
   8.750   0.9596   0.04873   0.04389  -0.0185   0.0107   1.0000
   9.000   0.9523   0.05383   0.04937  -0.0162   0.0105   1.0000
   9.250   0.9468   0.05804   0.05393  -0.0142   0.0103   1.0000
   9.500   0.9349   0.06241   0.05855  -0.0122   0.0102   1.0000
   9.750   0.9178   0.06673   0.06307  -0.0110   0.0101   1.0000
  10.000   0.8993   0.07171   0.06820  -0.0118   0.0102   1.0000
  10.250   0.8799   0.07745   0.07405  -0.0146   0.0104   1.0000
  10.500   0.7669   0.07182   0.06864  -0.0055   0.0113   1.0000
  10.750   0.7196   0.08863   0.08554  -0.0148   0.0122   1.0000
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