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=1,000,000 Ncrit=9


Details Polar file
Airfoil: PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il)
Reynolds number: 1,000,000
Max Cl/Cd: 88.39 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-pfcm-il-1000000.txt
Download as CSV file: xf-pfcm-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5944   0.08880   0.08725  -0.0088   1.0000   0.0079
  -9.000  -0.5978   0.08327   0.08173  -0.0126   1.0000   0.0080
  -8.750  -0.6011   0.07737   0.07586  -0.0175   1.0000   0.0080
  -8.500  -0.6049   0.06902   0.06751  -0.0286   1.0000   0.0079
  -8.250  -0.6062   0.06200   0.06040  -0.0343   1.0000   0.0080
  -8.000  -0.6049   0.05590   0.05417  -0.0377   1.0000   0.0081
  -7.750  -0.5988   0.05116   0.04929  -0.0395   1.0000   0.0082
  -7.500  -0.5888   0.04749   0.04547  -0.0404   1.0000   0.0084
  -7.250  -0.5769   0.04411   0.04195  -0.0408   1.0000   0.0087
  -7.000  -0.5644   0.04104   0.03874  -0.0406   1.0000   0.0092
  -6.750  -0.5525   0.03774   0.03526  -0.0395   1.0000   0.0102
  -6.000  -0.4935   0.01875   0.01464  -0.0390   0.9953   0.0093
  -5.750  -0.4616   0.01702   0.01269  -0.0402   0.9936   0.0101
  -5.500  -0.4310   0.01472   0.01008  -0.0409   0.9910   0.0106
  -5.250  -0.3986   0.01358   0.00876  -0.0420   0.9884   0.0113
  -5.000  -0.3658   0.01300   0.00807  -0.0432   0.9855   0.0118
  -4.750  -0.3343   0.01149   0.00640  -0.0442   0.9828   0.0122
  -4.500  -0.3078   0.00998   0.00477  -0.0442   0.9762   0.0133
  -4.250  -0.2782   0.00944   0.00419  -0.0447   0.9706   0.0144
  -4.000  -0.2512   0.00899   0.00369  -0.0446   0.9629   0.0155
  -3.750  -0.2245   0.00862   0.00325  -0.0443   0.9555   0.0165
  -3.500  -0.1984   0.00829   0.00286  -0.0439   0.9468   0.0173
  -3.250  -0.1722   0.00803   0.00255  -0.0435   0.9388   0.0180
  -3.000  -0.1460   0.00762   0.00203  -0.0431   0.9300   0.0211
  -2.750  -0.1191   0.00739   0.00176  -0.0428   0.9216   0.0265
  -2.500  -0.0925   0.00710   0.00152  -0.0425   0.9132   0.0515
  -2.250  -0.0659   0.00661   0.00133  -0.0426   0.9041   0.1438
  -2.000  -0.0402   0.00579   0.00112  -0.0427   0.8953   0.3363
  -1.750  -0.0146   0.00511   0.00101  -0.0426   0.8863   0.5240
  -1.500   0.0125   0.00487   0.00098  -0.0426   0.8769   0.5975
  -1.250   0.0398   0.00475   0.00096  -0.0425   0.8677   0.6448
  -1.000   0.0674   0.00470   0.00093  -0.0424   0.8580   0.6707
  -0.750   0.0948   0.00460   0.00094  -0.0423   0.8476   0.7149
  -0.500   0.1221   0.00454   0.00094  -0.0421   0.8372   0.7495
  -0.250   0.1498   0.00451   0.00093  -0.0420   0.8263   0.7700
   0.000   0.1774   0.00449   0.00092  -0.0420   0.8148   0.7878
   0.250   0.2051   0.00447   0.00091  -0.0419   0.8029   0.8035
   0.500   0.2326   0.00444   0.00091  -0.0418   0.7903   0.8195
   0.750   0.2597   0.00442   0.00092  -0.0416   0.7754   0.8401
   1.000   0.2859   0.00439   0.00094  -0.0411   0.7577   0.8697
   1.250   0.3098   0.00432   0.00097  -0.0401   0.7393   0.9181
   1.500   0.3401   0.00433   0.00098  -0.0404   0.7143   0.9764
   1.750   0.3740   0.00448   0.00098  -0.0419   0.6740   1.0000
   2.000   0.4008   0.00467   0.00102  -0.0417   0.6349   1.0000
   2.250   0.4276   0.00489   0.00110  -0.0416   0.5925   1.0000
   2.500   0.4543   0.00514   0.00118  -0.0416   0.5450   1.0000
   2.750   0.4802   0.00555   0.00131  -0.0414   0.4709   1.0000
   3.000   0.5023   0.00668   0.00166  -0.0409   0.2771   1.0000
   3.250   0.5255   0.00766   0.00203  -0.0406   0.1292   1.0000
   3.500   0.5506   0.00830   0.00235  -0.0404   0.0536   1.0000
   3.750   0.5765   0.00881   0.00273  -0.0402   0.0244   1.0000
   4.000   0.6036   0.00906   0.00300  -0.0401   0.0218   1.0000
   4.250   0.6303   0.00939   0.00333  -0.0400   0.0188   1.0000
   4.500   0.6551   0.01010   0.00417  -0.0395   0.0160   1.0000
   4.750   0.6815   0.01046   0.00456  -0.0393   0.0155   1.0000
   5.000   0.7074   0.01090   0.00505  -0.0390   0.0149   1.0000
   5.250   0.7328   0.01140   0.00560  -0.0387   0.0141   1.0000
   5.500   0.7585   0.01184   0.00607  -0.0384   0.0130   1.0000
   5.750   0.7841   0.01226   0.00652  -0.0382   0.0119   1.0000
   6.000   0.8065   0.01333   0.00767  -0.0373   0.0110   1.0000
   6.250   0.8258   0.01551   0.01003  -0.0359   0.0102   1.0000
   6.500   0.8516   0.01591   0.01049  -0.0356   0.0099   1.0000
   6.750   0.8764   0.01667   0.01134  -0.0351   0.0094   1.0000
   7.000   0.9010   0.01742   0.01218  -0.0347   0.0087   1.0000
   7.250   0.9258   0.01790   0.01274  -0.0344   0.0080   1.0000
   7.500   0.9500   0.01848   0.01338  -0.0341   0.0075   1.0000
   7.750   0.9731   0.01936   0.01434  -0.0336   0.0072   1.0000
   8.000   0.9919   0.02163   0.01684  -0.0324   0.0068   1.0000
   8.250   0.9975   0.02775   0.02365  -0.0293   0.0064   1.0000
   8.500   1.0187   0.02867   0.02475  -0.0286   0.0062   1.0000
   8.750   1.0321   0.03162   0.02805  -0.0268   0.0060   1.0000
   9.000   1.0187   0.04124   0.03841  -0.0219   0.0057   1.0000
   9.250   0.9959   0.05099   0.04868  -0.0177   0.0058   1.0000
   9.500   0.9760   0.05764   0.05561  -0.0152   0.0061   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)