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PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il)
Reynolds number: 200,000
Max Cl/Cd: 61.64 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-pfcm-il-200000.txt
Download as CSV file: xf-pfcm-il-200000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.4476   0.11842   0.11498  -0.0083   1.0000   0.0342
 -10.750  -0.4481   0.11471   0.11127  -0.0097   1.0000   0.0359
  -7.500  -0.5657   0.06440   0.06086  -0.0367   1.0000   0.0419
  -7.250  -0.5577   0.06078   0.05718  -0.0376   1.0000   0.0435
  -7.000  -0.5489   0.05684   0.05311  -0.0388   1.0000   0.0455
  -6.750  -0.5383   0.05280   0.04884  -0.0399   1.0000   0.0486
  -6.250  -0.4965   0.03175   0.02770  -0.0382   1.0000   0.0548
  -6.000  -0.4874   0.02903   0.02492  -0.0368   1.0000   0.0574
  -5.750  -0.4760   0.02950   0.02465  -0.0339   1.0000   0.0651
  -5.500  -0.4702   0.02321   0.01845  -0.0338   1.0000   0.0677
  -5.250  -0.4577   0.02105   0.01628  -0.0324   1.0000   0.0718
  -5.000  -0.4456   0.01876   0.01364  -0.0310   1.0000   0.0818
  -4.750  -0.4320   0.01702   0.01173  -0.0298   1.0000   0.0950
  -4.500  -0.4173   0.01523   0.00988  -0.0287   1.0000   0.1097
  -4.250  -0.3703   0.02206   0.01464  -0.0284   1.0000   0.0393
  -4.000  -0.3469   0.01957   0.01192  -0.0275   1.0000   0.0385
  -3.750  -0.3233   0.01815   0.01030  -0.0266   1.0000   0.0388
  -3.500  -0.2989   0.01622   0.00828  -0.0262   1.0000   0.0423
  -3.250  -0.2748   0.01524   0.00728  -0.0256   1.0000   0.0444
  -3.000  -0.2504   0.01441   0.00643  -0.0251   1.0000   0.0471
  -2.750  -0.2252   0.01378   0.00574  -0.0248   1.0000   0.0513
  -2.500  -0.1990   0.01304   0.00502  -0.0250   1.0000   0.0635
  -2.250  -0.1649   0.01004   0.00421  -0.0278   1.0000   0.5743
  -2.000  -0.1337   0.00971   0.00441  -0.0281   0.9964   0.7367
  -1.750  -0.0987   0.00959   0.00441  -0.0292   0.9907   0.8014
  -1.500  -0.0695   0.00942   0.00447  -0.0285   0.9846   0.8746
  -1.250  -0.0335   0.00940   0.00452  -0.0289   0.9804   0.9589
  -1.000   0.0204   0.00938   0.00439  -0.0348   0.9765   0.9954
  -0.750   0.0672   0.00933   0.00423  -0.0396   0.9712   1.0000
  -0.500   0.1093   0.00929   0.00410  -0.0432   0.9644   1.0000
  -0.250   0.1525   0.00925   0.00398  -0.0469   0.9582   1.0000
   0.000   0.1916   0.00922   0.00391  -0.0496   0.9509   1.0000
   0.250   0.2301   0.00917   0.00384  -0.0520   0.9434   1.0000
   0.500   0.2629   0.00916   0.00381  -0.0532   0.9335   1.0000
   0.750   0.2958   0.00914   0.00378  -0.0542   0.9248   1.0000
   1.000   0.3254   0.00913   0.00377  -0.0545   0.9147   1.0000
   1.250   0.3522   0.00916   0.00383  -0.0542   0.9031   1.0000
   1.500   0.3786   0.00920   0.00388  -0.0538   0.8918   1.0000
   1.750   0.4047   0.00923   0.00394  -0.0532   0.8804   1.0000
   2.000   0.4301   0.00925   0.00398  -0.0523   0.8679   1.0000
   2.250   0.4545   0.00924   0.00403  -0.0511   0.8524   1.0000
   2.500   0.4762   0.00908   0.00383  -0.0488   0.8274   1.0000
   2.750   0.4973   0.00890   0.00357  -0.0462   0.7910   1.0000
   3.000   0.5206   0.00889   0.00350  -0.0446   0.7553   1.0000
   3.250   0.5441   0.00897   0.00353  -0.0431   0.7126   1.0000
   3.500   0.5665   0.00919   0.00353  -0.0413   0.6437   1.0000
   3.750   0.5836   0.01002   0.00361  -0.0388   0.4688   1.0000
   4.250   0.6111   0.01459   0.00602  -0.0351   0.0551   1.0000
   4.500   0.6323   0.01584   0.00722  -0.0341   0.0468   1.0000
   4.750   0.6559   0.01680   0.00830  -0.0332   0.0436   1.0000
   5.000   0.6795   0.01800   0.00955  -0.0323   0.0410   1.0000
   5.250   0.7039   0.01934   0.01092  -0.0315   0.0388   1.0000
   5.500   0.7277   0.02203   0.01362  -0.0309   0.0353   1.0000
   5.750   0.7537   0.02392   0.01572  -0.0302   0.0347   1.0000
   6.000   0.7791   0.02649   0.01857  -0.0294   0.0347   1.0000
   6.250   0.8045   0.02912   0.02149  -0.0286   0.0358   1.0000
   6.500   0.8286   0.03131   0.02402  -0.0274   0.0363   1.0000
   6.750   0.8562   0.04054   0.03473  -0.0222   0.0758   1.0000
   7.000   0.8680   0.04366   0.03830  -0.0206   0.0668   1.0000
   7.250   0.8843   0.04661   0.04135  -0.0198   0.0629   1.0000
   7.750   0.8989   0.05423   0.04971  -0.0167   0.0538   1.0000
   8.000   0.9069   0.05765   0.05332  -0.0157   0.0512   1.0000
   8.250   0.9163   0.06117   0.05690  -0.0150   0.0494   1.0000
   8.500   0.9257   0.06820   0.06375  -0.0155   0.0477   1.0000
  10.750   0.8050   0.11924   0.11592  -0.0393   0.0434   1.0000
  11.000   0.8019   0.12482   0.12147  -0.0424   0.0421   1.0000
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