Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il) Xfoil prediction polar at RE=500,000 Ncrit=0


Details Polar file
Airfoil: PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il)
Reynolds number: 500,000
Max Cl/Cd: 78.74 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-pfcm-il-500000.txt
Download as CSV file: xf-pfcm-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.4537   0.11976   0.11758  -0.0086   1.0000   0.0127
 -11.250  -0.4569   0.11497   0.11280  -0.0096   1.0000   0.0130
 -11.000  -0.4555   0.11138   0.10921  -0.0101   1.0000   0.0133
 -10.750  -0.4536   0.10784   0.10568  -0.0108   1.0000   0.0136
  -9.750  -0.5782   0.10190   0.09968  -0.0055   1.0000   0.0129
  -9.500  -0.5797   0.09704   0.09483  -0.0078   1.0000   0.0131
  -9.250  -0.5786   0.09287   0.09068  -0.0098   1.0000   0.0133
  -5.750  -0.4669   0.01503   0.01123  -0.0363   0.9962   0.0226
  -5.500  -0.4376   0.01288   0.00895  -0.0383   0.9940   0.0246
  -5.000  -0.4114   0.01906   0.01376  -0.0350   0.9955   0.0187
  -4.750  -0.3774   0.01709   0.01148  -0.0363   0.9934   0.0194
  -4.500  -0.3434   0.01556   0.00969  -0.0376   0.9908   0.0211
  -4.250  -0.3102   0.01342   0.00736  -0.0388   0.9887   0.0218
  -4.000  -0.2755   0.01211   0.00597  -0.0405   0.9869   0.0227
  -3.750  -0.2414   0.01133   0.00513  -0.0420   0.9839   0.0244
  -3.500  -0.2067   0.01058   0.00433  -0.0436   0.9808   0.0256
  -3.250  -0.1712   0.00994   0.00361  -0.0454   0.9779   0.0273
  -3.000  -0.1374   0.00948   0.00309  -0.0467   0.9737   0.0308
  -2.750  -0.1052   0.00888   0.00246  -0.0477   0.9682   0.0474
  -2.500  -0.0760   0.00769   0.00206  -0.0488   0.9624   0.2533
  -2.250  -0.0508   0.00648   0.00189  -0.0490   0.9548   0.5506
  -2.000  -0.0245   0.00619   0.00188  -0.0486   0.9466   0.6406
  -1.750   0.0018   0.00605   0.00184  -0.0480   0.9387   0.6900
  -1.500   0.0276   0.00598   0.00179  -0.0474   0.9288   0.7162
  -1.250   0.0528   0.00589   0.00178  -0.0466   0.9195   0.7533
  -1.000   0.0770   0.00578   0.00177  -0.0454   0.9106   0.7946
  -0.750   0.1022   0.00568   0.00172  -0.0446   0.9002   0.8181
  -0.500   0.1278   0.00559   0.00167  -0.0439   0.8900   0.8396
  -0.250   0.1529   0.00551   0.00163  -0.0431   0.8803   0.8617
   0.000   0.1773   0.00543   0.00161  -0.0421   0.8702   0.8901
   0.250   0.2025   0.00535   0.00160  -0.0411   0.8591   0.9288
   0.500   0.2374   0.00532   0.00157  -0.0425   0.8492   0.9692
   0.750   0.2775   0.00532   0.00154  -0.0453   0.8396   0.9993
   1.000   0.3046   0.00537   0.00153  -0.0452   0.8281   1.0000
   1.250   0.3315   0.00541   0.00154  -0.0451   0.8149   1.0000
   1.500   0.3582   0.00547   0.00154  -0.0448   0.7987   1.0000
   1.750   0.3848   0.00553   0.00155  -0.0445   0.7782   1.0000
   2.000   0.4103   0.00565   0.00153  -0.0438   0.7464   1.0000
   2.250   0.4360   0.00580   0.00156  -0.0433   0.7081   1.0000
   2.500   0.4620   0.00598   0.00161  -0.0429   0.6700   1.0000
   2.750   0.4878   0.00622   0.00169  -0.0424   0.6235   1.0000
   3.000   0.5134   0.00652   0.00180  -0.0420   0.5699   1.0000
   3.250   0.5374   0.00705   0.00199  -0.0415   0.4763   1.0000
   3.500   0.5560   0.00860   0.00250  -0.0406   0.2380   1.0000
   3.750   0.5760   0.01009   0.00315  -0.0399   0.0570   1.0000
   4.000   0.6012   0.01072   0.00370  -0.0395   0.0344   1.0000
   4.250   0.6261   0.01137   0.00438  -0.0390   0.0289   1.0000
   4.500   0.6506   0.01208   0.00520  -0.0384   0.0272   1.0000
   4.750   0.6756   0.01265   0.00584  -0.0380   0.0259   1.0000
   5.000   0.7001   0.01335   0.00661  -0.0374   0.0243   1.0000
   5.250   0.7247   0.01400   0.00731  -0.0369   0.0222   1.0000
   5.500   0.7485   0.01487   0.00823  -0.0362   0.0209   1.0000
   5.750   0.7719   0.01593   0.00939  -0.0354   0.0197   1.0000
   6.000   0.7915   0.01952   0.01318  -0.0340   0.0175   1.0000
   6.250   0.8176   0.01958   0.01334  -0.0337   0.0166   1.0000
   6.500   0.8421   0.02105   0.01500  -0.0329   0.0157   1.0000
   6.750   0.8654   0.02319   0.01741  -0.0319   0.0149   1.0000
   7.000   0.8866   0.02616   0.02074  -0.0305   0.0144   1.0000
   7.250   0.9081   0.02759   0.02237  -0.0297   0.0131   1.0000
   7.500   0.9298   0.02800   0.02279  -0.0295   0.0119   1.0000
   7.750   0.9457   0.03111   0.02627  -0.0279   0.0116   1.0000
<< Back to PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il)

Polar data table (+)

Polar graphs


<< Back to PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il)