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Coanda 1 (coanda1-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: Coanda 1 (coanda1-il)
Reynolds number: 50,000
Max Cl/Cd: 37.96 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-coanda1-il-50000.txt
Download as CSV file: xf-coanda1-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Coanda 1                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4394   0.10940   0.10251  -0.0132   1.0000   0.1151
  -8.500  -0.4475   0.10808   0.10131  -0.0150   1.0000   0.1165
  -8.250  -0.4585   0.10708   0.10044  -0.0165   1.0000   0.1170
  -8.000  -0.4418   0.10055   0.09393  -0.0147   1.0000   0.1196
  -7.750  -0.4373   0.09695   0.09037  -0.0139   1.0000   0.1231
  -7.500  -0.4391   0.09436   0.08786  -0.0142   1.0000   0.1265
  -7.250  -0.4437   0.09260   0.08618  -0.0168   1.0000   0.1293
  -7.000  -0.4461   0.09078   0.08441  -0.0203   1.0000   0.1306
  -6.750  -0.4338   0.08495   0.07864  -0.0158   1.0000   0.1354
  -6.500  -0.3724   0.07028   0.06450  -0.0162   1.0000   0.1509
  -6.250  -0.3794   0.06795   0.06221  -0.0186   1.0000   0.1570
  -6.000  -0.4206   0.07540   0.06918  -0.0168   1.0000   0.1523
  -5.750  -0.4158   0.07241   0.06618  -0.0183   1.0000   0.1599
  -5.500  -0.4099   0.06931   0.06307  -0.0178   1.0000   0.1735
  -5.250  -0.4037   0.06603   0.05984  -0.0155   1.0000   0.1908
  -5.000  -0.1647   0.04666   0.03995   0.0039   1.0000   0.9219
  -4.750  -0.2066   0.04705   0.04058   0.0131   1.0000   0.8794
  -4.500  -0.2434   0.04692   0.04068   0.0202   1.0000   0.8399
  -4.250  -0.2725   0.04623   0.04019   0.0263   1.0000   0.8226
  -4.000  -0.2614   0.04290   0.03690   0.0259   1.0000   0.8344
  -3.750  -0.3984   0.04912   0.04343   0.0113   1.0000   0.4512
  -3.500  -0.4035   0.04633   0.04081   0.0203   1.0000   0.5152
  -3.250  -0.4073   0.04349   0.03811   0.0284   1.0000   0.5712
  -3.000  -0.4152   0.04033   0.03513   0.0392   1.0000   0.6395
  -2.750  -0.2486   0.03329   0.02463  -0.0095   1.0000   0.2091
  -2.500  -0.2375   0.03099   0.02227  -0.0067   1.0000   0.2608
  -2.250  -0.2164   0.02916   0.01998  -0.0052   1.0000   0.2721
  -2.000  -0.1911   0.02778   0.01809  -0.0045   1.0000   0.2631
  -1.750  -0.1656   0.02678   0.01655  -0.0038   1.0000   0.2561
  -1.500  -0.1412   0.02597   0.01532  -0.0030   1.0000   0.2539
  -1.250  -0.1173   0.02517   0.01430  -0.0024   1.0000   0.2565
  -1.000  -0.0930   0.02462   0.01347  -0.0018   1.0000   0.2556
  -0.750  -0.0665   0.02411   0.01273  -0.0017   1.0000   0.2534
  -0.500  -0.0345   0.02368   0.01206  -0.0027   1.0000   0.2521
  -0.250  -0.0022   0.02336   0.01149  -0.0039   1.0000   0.2517
   0.000   0.0281   0.02315   0.01114  -0.0048   1.0000   0.2523
   0.250   0.0560   0.02302   0.01092  -0.0053   1.0000   0.2544
   0.500   0.0817   0.02293   0.01086  -0.0056   1.0000   0.2568
   0.750   0.1050   0.02294   0.01093  -0.0057   1.0000   0.2646
   1.000   0.1265   0.02304   0.01110  -0.0056   1.0000   0.2753
   1.250   0.1467   0.02326   0.01141  -0.0054   1.0000   0.2852
   1.500   0.1660   0.02357   0.01184  -0.0052   1.0000   0.2966
   1.750   0.1854   0.02389   0.01241  -0.0052   1.0000   0.3178
   2.000   0.2701   0.02345   0.01355  -0.0178   1.0000   1.0000
   2.250   0.3338   0.02459   0.01455  -0.0266   0.9791   1.0000
   2.500   0.4304   0.02501   0.01498  -0.0401   0.9415   1.0000
   2.750   0.5161   0.02450   0.01469  -0.0502   0.9033   1.0000
   3.000   0.5940   0.02352   0.01401  -0.0579   0.8669   1.0000
   3.250   0.6616   0.02215   0.01304  -0.0626   0.8257   1.0000
   3.500   0.7149   0.02093   0.01205  -0.0640   0.7739   1.0000
   3.750   0.7542   0.02037   0.01146  -0.0628   0.7074   1.0000
   4.000   0.7779   0.02055   0.01136  -0.0593   0.6294   1.0000
   4.250   0.7910   0.02084   0.01120  -0.0545   0.5514   1.0000
   4.500   0.8052   0.02142   0.01149  -0.0509   0.4853   1.0000
   4.750   0.8256   0.02222   0.01217  -0.0490   0.4387   1.0000
   5.000   0.8469   0.02301   0.01290  -0.0473   0.4030   1.0000
   5.250   0.8664   0.02373   0.01363  -0.0454   0.3698   1.0000
   5.500   0.8806   0.02432   0.01414  -0.0427   0.3307   1.0000
   5.750   0.8989   0.02497   0.01496  -0.0407   0.3027   1.0000
   6.000   0.9070   0.02544   0.01539  -0.0371   0.2471   1.0000
   6.250   0.9167   0.02659   0.01631  -0.0340   0.1603   1.0000
   6.500   0.9303   0.02866   0.01796  -0.0314   0.0998   1.0000
   6.750   0.9451   0.03057   0.01972  -0.0289   0.0843   1.0000
   7.000   0.9607   0.03246   0.02164  -0.0266   0.0762   1.0000
   7.250   0.9758   0.03453   0.02367  -0.0244   0.0695   1.0000
   7.500   0.9966   0.03681   0.02619  -0.0226   0.0661   1.0000
   7.750   1.0198   0.03959   0.02929  -0.0211   0.0647   1.0000
   8.000   1.0405   0.04287   0.03298  -0.0194   0.0643   1.0000
   8.250   1.0553   0.04660   0.03722  -0.0173   0.0646   1.0000
   8.500   1.0640   0.05061   0.04172  -0.0148   0.0653   1.0000
   8.750   1.0677   0.05476   0.04633  -0.0121   0.0662   1.0000
   9.000   1.0669   0.05914   0.05109  -0.0093   0.0672   1.0000
   9.250   1.0639   0.06357   0.05580  -0.0068   0.0681   1.0000
   9.500   1.0623   0.06823   0.06064  -0.0048   0.0689   1.0000
   9.750   1.0392   0.07184   0.06475  -0.0014   0.0702   1.0000
  10.000   1.0063   0.07597   0.06914   0.0016   0.0709   1.0000
  10.250   0.9728   0.08085   0.07419   0.0022   0.0716   1.0000
  10.500   0.9387   0.08751   0.08095  -0.0009   0.0725   1.0000
  10.750   0.9048   0.09700   0.09045  -0.0076   0.0738   1.0000
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