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

USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: USA 41 AIRFOIL (usa41-il)
Reynolds number: 1,000,000
Max Cl/Cd: 129.19 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa41-il-1000000.txt
Download as CSV file: xf-usa41-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 41 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3487   0.08614   0.08460  -0.0235   1.0000   0.0122
  -7.500  -0.3539   0.08404   0.08254  -0.0222   1.0000   0.0122
  -7.250  -0.3551   0.08206   0.08060  -0.0218   0.9994   0.0130
  -7.000  -0.3210   0.07486   0.07338  -0.0357   0.9950   0.0147
  -6.750  -0.2948   0.06923   0.06774  -0.0442   0.9903   0.0148
  -6.500  -0.2650   0.06343   0.06190  -0.0533   0.9857   0.0148
  -6.250  -0.2347   0.05678   0.05520  -0.0631   0.9797   0.0148
  -6.000  -0.1854   0.01919   0.01619  -0.0994   0.9658   0.0151
  -5.750  -0.1542   0.01818   0.01507  -0.1010   0.9603   0.0157
  -5.500  -0.1208   0.01842   0.01535  -0.1023   0.9554   0.0162
  -5.250  -0.0903   0.01789   0.01473  -0.1032   0.9470   0.0171
  -5.000  -0.0606   0.01630   0.01281  -0.1040   0.9378   0.0186
  -4.500  -0.0076   0.01332   0.00922  -0.1041   0.9171   0.0210
  -4.250   0.0189   0.01302   0.00889  -0.1040   0.9071   0.0219
  -4.000   0.0454   0.01271   0.00849  -0.1037   0.8975   0.0232
  -3.750   0.0714   0.01215   0.00777  -0.1032   0.8873   0.0244
  -3.500   0.0975   0.01151   0.00696  -0.1026   0.8775   0.0251
  -3.250   0.1241   0.01120   0.00651  -0.1022   0.8677   0.0259
  -3.000   0.1491   0.01001   0.00508  -0.1016   0.8570   0.0278
  -2.750   0.1748   0.00935   0.00436  -0.1011   0.8455   0.0288
  -2.500   0.2009   0.00887   0.00378  -0.1005   0.8329   0.0292
  -2.250   0.2269   0.00844   0.00326  -0.1000   0.8187   0.0294
  -2.000   0.2528   0.00808   0.00281  -0.0993   0.8021   0.0297
  -1.500   0.3043   0.00761   0.00214  -0.0981   0.7634   0.0314
  -1.250   0.3298   0.00745   0.00187  -0.0974   0.7416   0.0322
  -1.000   0.3555   0.00733   0.00165  -0.0968   0.7196   0.0329
  -0.750   0.3812   0.00727   0.00147  -0.0962   0.6974   0.0334
  -0.500   0.4071   0.00722   0.00132  -0.0956   0.6741   0.0338
  -0.250   0.4329   0.00721   0.00120  -0.0950   0.6500   0.0341
   0.000   0.4588   0.00722   0.00110  -0.0945   0.6264   0.0345
   0.250   0.4848   0.00726   0.00104  -0.0940   0.6030   0.0352
   0.750   0.5370   0.00734   0.00092  -0.0930   0.5628   0.0381
   1.000   0.5634   0.00738   0.00091  -0.0926   0.5459   0.0450
   1.250   0.5890   0.00722   0.00099  -0.0922   0.5310   0.1575
   1.500   0.6155   0.00729   0.00107  -0.0918   0.5175   0.1836
   1.750   0.6420   0.00737   0.00114  -0.0915   0.5043   0.1976
   2.000   0.6684   0.00745   0.00120  -0.0911   0.4907   0.2106
   2.250   0.6948   0.00752   0.00126  -0.0908   0.4770   0.2221
   2.750   0.7468   0.00751   0.00142  -0.0900   0.4502   0.3231
   3.000   0.7900   0.00633   0.00167  -0.0940   0.4352   1.0000
   3.250   0.8163   0.00646   0.00175  -0.0936   0.4231   1.0000
   3.500   0.8424   0.00659   0.00186  -0.0932   0.4120   1.0000
   3.750   0.8684   0.00674   0.00196  -0.0928   0.3998   1.0000
   4.000   0.8940   0.00692   0.00208  -0.0923   0.3830   1.0000
   4.250   0.9189   0.00716   0.00221  -0.0917   0.3598   1.0000
   4.500   0.9435   0.00744   0.00238  -0.0911   0.3336   1.0000
   4.750   0.9688   0.00765   0.00254  -0.0906   0.3173   1.0000
   5.000   0.9939   0.00788   0.00271  -0.0901   0.2985   1.0000
   5.250   1.0181   0.00820   0.00292  -0.0894   0.2717   1.0000
   5.500   1.0408   0.00867   0.00320  -0.0885   0.2301   1.0000
   5.750   1.0603   0.00950   0.00366  -0.0872   0.1626   1.0000
   6.000   1.0749   0.01091   0.00450  -0.0852   0.0612   1.0000
   6.250   1.0944   0.01180   0.00516  -0.0838   0.0227   1.0000
   6.500   1.1180   0.01219   0.00558  -0.0830   0.0187   1.0000
   6.750   1.1409   0.01265   0.00606  -0.0821   0.0162   1.0000
   7.000   1.1629   0.01323   0.00671  -0.0810   0.0144   1.0000
   7.250   1.1860   0.01363   0.00715  -0.0802   0.0135   1.0000
   7.500   1.2084   0.01409   0.00765  -0.0793   0.0126   1.0000
   7.750   1.2302   0.01460   0.00820  -0.0783   0.0116   1.0000
   8.000   1.2462   0.01571   0.00943  -0.0763   0.0104   1.0000
   8.250   1.2669   0.01628   0.01006  -0.0751   0.0100   1.0000
   8.500   1.2870   0.01687   0.01071  -0.0739   0.0096   1.0000
   8.750   1.3060   0.01754   0.01145  -0.0725   0.0092   1.0000
   9.000   1.3243   0.01823   0.01220  -0.0710   0.0088   1.0000
   9.250   1.3411   0.01903   0.01308  -0.0693   0.0085   1.0000
   9.500   1.3586   0.01971   0.01380  -0.0678   0.0080   1.0000
   9.750   1.3740   0.02053   0.01466  -0.0660   0.0077   1.0000
  10.000   1.3783   0.02224   0.01648  -0.0626   0.0072   1.0000
  10.250   1.3843   0.02365   0.01802  -0.0593   0.0069   1.0000
  10.500   1.3953   0.02442   0.01886  -0.0568   0.0068   1.0000
  10.750   1.4034   0.02544   0.01999  -0.0540   0.0066   1.0000
  11.000   1.4112   0.02651   0.02115  -0.0513   0.0064   1.0000
  11.250   1.4168   0.02784   0.02259  -0.0485   0.0063   1.0000
  11.500   1.4206   0.02943   0.02430  -0.0458   0.0061   1.0000
  11.750   1.4247   0.03106   0.02605  -0.0435   0.0061   1.0000
  12.000   1.4297   0.03253   0.02762  -0.0414   0.0058   1.0000
  12.250   1.4338   0.03417   0.02936  -0.0396   0.0057   1.0000
  12.500   1.4337   0.03646   0.03179  -0.0376   0.0056   1.0000
  12.750   1.4355   0.03850   0.03395  -0.0362   0.0056   1.0000
  13.000   1.4394   0.04021   0.03574  -0.0353   0.0054   1.0000
  13.250   1.4319   0.04378   0.03953  -0.0338   0.0054   1.0000
  13.500   1.4320   0.04621   0.04204  -0.0334   0.0053   1.0000
  13.750   1.4332   0.04863   0.04454  -0.0334   0.0051   1.0000
  14.000   1.4258   0.05247   0.04854  -0.0333   0.0051   1.0000
  14.250   1.4192   0.05634   0.05254  -0.0338   0.0051   1.0000
  14.500   1.4133   0.06021   0.05650  -0.0347   0.0049   1.0000
  14.750   1.4021   0.06516   0.06163  -0.0359   0.0050   1.0000
  15.000   1.3892   0.07061   0.06724  -0.0377   0.0049   1.0000
  15.250   1.3731   0.07683   0.07362  -0.0400   0.0049   1.0000
  15.500   1.3570   0.08336   0.08031  -0.0428   0.0048   1.0000
  15.750   1.3390   0.09058   0.08770  -0.0462   0.0048   1.0000
  16.000   1.3270   0.09702   0.09428  -0.0494   0.0049   1.0000
  16.250   1.3090   0.10490   0.10232  -0.0536   0.0049   1.0000
  16.500   1.2901   0.11337   0.11094  -0.0583   0.0049   1.0000
  16.750   1.2691   0.12265   0.12037  -0.0636   0.0048   1.0000
  17.000   1.2571   0.13037   0.12823  -0.0682   0.0049   1.0000
  17.250   1.2364   0.14060   0.13862  -0.0743   0.0050   1.0000
  17.500   1.2176   0.15076   0.14890  -0.0806   0.0049   1.0000
<< Back to USA 41 AIRFOIL (usa41-il)

Polar data table (+)

Polar graphs


<< Back to USA 41 AIRFOIL (usa41-il)