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USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 100,000
Max Cl/Cd: 51.82 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa46-il-100000-n5.txt
Download as CSV file: xf-usa46-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 46 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5078   0.08777   0.08288  -0.0154   1.0000   0.0229
  -8.250  -0.5088   0.08363   0.07881  -0.0173   1.0000   0.0228
  -8.000  -0.5121   0.07932   0.07459  -0.0193   1.0000   0.0228
  -7.750  -0.5131   0.07440   0.06973  -0.0229   1.0000   0.0227
  -7.500  -0.5110   0.06862   0.06400  -0.0284   1.0000   0.0226
  -7.250  -0.5061   0.06180   0.05718  -0.0352   1.0000   0.0224
  -7.000  -0.4994   0.05325   0.04851  -0.0426   1.0000   0.0223
  -6.750  -0.4910   0.04302   0.03786  -0.0498   1.0000   0.0222
  -6.500  -0.4772   0.03459   0.02863  -0.0542   1.0000   0.0225
  -6.250  -0.4576   0.02909   0.02224  -0.0557   1.0000   0.0235
  -6.000  -0.4362   0.02697   0.01991  -0.0559   1.0000   0.0260
  -5.750  -0.4131   0.02488   0.01742  -0.0557   1.0000   0.0294
  -5.500  -0.3890   0.02221   0.01412  -0.0553   1.0000   0.0330
  -5.250  -0.3655   0.02119   0.01304  -0.0550   1.0000   0.0395
  -5.000  -0.3414   0.01972   0.01134  -0.0544   1.0000   0.0475
  -4.750  -0.3171   0.01880   0.01017  -0.0540   1.0000   0.0592
  -4.500  -0.2929   0.01813   0.00932  -0.0535   1.0000   0.0718
  -4.250  -0.2687   0.01734   0.00842  -0.0528   1.0000   0.0798
  -4.000  -0.2447   0.01670   0.00771  -0.0522   1.0000   0.0888
  -3.750  -0.2206   0.01613   0.00704  -0.0515   1.0000   0.0983
  -3.500  -0.1966   0.01561   0.00645  -0.0508   1.0000   0.1072
  -3.250  -0.1727   0.01516   0.00598  -0.0501   1.0000   0.1183
  -3.000  -0.1489   0.01480   0.00559  -0.0494   1.0000   0.1338
  -2.750  -0.1252   0.01453   0.00527  -0.0487   1.0000   0.1535
  -2.500  -0.1019   0.01422   0.00495  -0.0480   1.0000   0.1710
  -2.250  -0.0789   0.01396   0.00471  -0.0473   1.0000   0.1906
  -2.000  -0.0559   0.01375   0.00450  -0.0466   1.0000   0.2100
  -1.750  -0.0305   0.01350   0.00433  -0.0465   0.9989   0.2325
  -1.500   0.0066   0.01319   0.00417  -0.0488   0.9904   0.2700
  -1.250   0.0427   0.01273   0.00406  -0.0509   0.9819   0.3479
  -1.000   0.0769   0.01196   0.00410  -0.0524   0.9740   0.5665
  -0.750   0.1085   0.01152   0.00416  -0.0525   0.9648   0.7488
  -0.500   0.1448   0.01112   0.00399  -0.0536   0.9546   0.8597
  -0.250   0.1942   0.01091   0.00377  -0.0578   0.9417   1.0000
   0.000   0.2351   0.01089   0.00363  -0.0602   0.9259   1.0000
   0.250   0.2733   0.01086   0.00352  -0.0620   0.9063   1.0000
   0.500   0.3102   0.01085   0.00343  -0.0634   0.8858   1.0000
   0.750   0.3436   0.01088   0.00338  -0.0642   0.8649   1.0000
   1.000   0.3752   0.01093   0.00339  -0.0646   0.8442   1.0000
   1.250   0.4050   0.01100   0.00342  -0.0646   0.8218   1.0000
   1.500   0.4343   0.01110   0.00347  -0.0645   0.7989   1.0000
   1.750   0.4622   0.01121   0.00357  -0.0641   0.7739   1.0000
   2.000   0.4897   0.01134   0.00367  -0.0636   0.7480   1.0000
   2.250   0.5169   0.01149   0.00379  -0.0631   0.7215   1.0000
   2.500   0.5436   0.01167   0.00397  -0.0624   0.6942   1.0000
   2.750   0.5700   0.01188   0.00415  -0.0618   0.6662   1.0000
   3.000   0.5961   0.01212   0.00436  -0.0611   0.6373   1.0000
   3.250   0.6218   0.01238   0.00464  -0.0603   0.6077   1.0000
   3.500   0.6472   0.01269   0.00492  -0.0595   0.5757   1.0000
   3.750   0.6720   0.01305   0.00524  -0.0586   0.5412   1.0000
   4.000   0.6965   0.01344   0.00560  -0.0576   0.5019   1.0000
   4.250   0.7184   0.01396   0.00588  -0.0563   0.4284   1.0000
   4.500   0.7390   0.01474   0.00626  -0.0550   0.3435   1.0000
   4.750   0.7586   0.01586   0.00682  -0.0538   0.2328   1.0000
   5.000   0.7740   0.01810   0.00794  -0.0525   0.0588   1.0000
   5.250   0.7953   0.01954   0.00930  -0.0514   0.0357   1.0000
   5.500   0.8173   0.02073   0.01060  -0.0503   0.0295   1.0000
   5.750   0.8391   0.02191   0.01206  -0.0492   0.0266   1.0000
   6.000   0.8599   0.02317   0.01344  -0.0481   0.0233   1.0000
   6.250   0.8778   0.02502   0.01536  -0.0466   0.0210   1.0000
   6.500   0.8988   0.02670   0.01719  -0.0453   0.0202   1.0000
   6.750   0.9208   0.02866   0.01933  -0.0442   0.0196   1.0000
   7.000   0.9434   0.03091   0.02182  -0.0430   0.0190   1.0000
   7.250   0.9653   0.03321   0.02445  -0.0419   0.0178   1.0000
   7.500   0.9853   0.03567   0.02728  -0.0407   0.0165   1.0000
   7.750   1.0030   0.03865   0.03068  -0.0393   0.0161   1.0000
   8.000   1.0176   0.04203   0.03464  -0.0377   0.0159   1.0000
   8.250   1.0281   0.04577   0.03887  -0.0360   0.0159   1.0000
   8.500   1.0345   0.04973   0.04331  -0.0343   0.0159   1.0000
   8.750   1.0362   0.05393   0.04795  -0.0326   0.0160   1.0000
   9.000   1.0331   0.05818   0.05259  -0.0311   0.0162   1.0000
   9.250   1.0247   0.06247   0.05720  -0.0298   0.0163   1.0000
   9.500   1.0101   0.06652   0.06149  -0.0284   0.0165   1.0000
   9.750   0.9918   0.07074   0.06589  -0.0281   0.0166   1.0000
  10.000   0.9731   0.07576   0.07107  -0.0301   0.0167   1.0000
  10.250   0.9529   0.08237   0.07781  -0.0347   0.0168   1.0000
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