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lrn1007 (lrn1007-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: lrn1007 (lrn1007-il)
Reynolds number: 200,000
Max Cl/Cd: 81.74 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-lrn1007-il-200000.txt
Download as CSV file: xf-lrn1007-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: lrn1007                                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.1928   0.08883   0.08529  -0.0680   0.9456   0.0044
  -7.750  -0.1814   0.08602   0.08248  -0.0706   0.9432   0.0062
  -7.500  -0.1716   0.08358   0.08005  -0.0728   0.9404   0.0069
  -7.250  -0.1613   0.08218   0.07866  -0.0753   0.9372   0.0082
  -7.000  -0.1483   0.08054   0.07703  -0.0792   0.9343   0.0085
  -6.750  -0.1322   0.07868   0.07515  -0.0838   0.9315   0.0087
  -6.250  -0.0954   0.07409   0.07051  -0.0930   0.9277   0.0089
  -6.000  -0.0754   0.07171   0.06809  -0.0970   0.9258   0.0090
  -5.750  -0.0557   0.06920   0.06552  -0.1001   0.9237   0.0091
  -5.500  -0.0356   0.06658   0.06284  -0.1027   0.9215   0.0091
  -5.250  -0.0148   0.06394   0.06013  -0.1051   0.9193   0.0092
  -5.000   0.0061   0.06115   0.05725  -0.1071   0.9176   0.0092
  -4.750   0.0279   0.05847   0.05448  -0.1090   0.9163   0.0092
  -4.500   0.0432   0.05451   0.05051  -0.1105   0.9142   0.0095
  -4.250   0.0538   0.05008   0.04609  -0.1113   0.9122   0.0100
  -4.000   0.0774   0.04828   0.04416  -0.1125   0.9104   0.0098
  -3.750   0.0903   0.04456   0.04038  -0.1131   0.9087   0.0106
  -3.500   0.1114   0.04189   0.03760  -0.1141   0.9068   0.0115
  -3.250   0.1364   0.03949   0.03506  -0.1152   0.9052   0.0126
  -3.000   0.1637   0.03720   0.03244  -0.1162   0.9039   0.0157
  -2.750   0.1826   0.03584   0.03094  -0.1154   0.9015   0.0171
  -2.500   0.2022   0.03499   0.02992  -0.1139   0.8984   0.0187
  -2.250   0.2299   0.03548   0.03011  -0.1125   0.8962   0.0200
  -2.000   0.2563   0.03479   0.02914  -0.1119   0.8942   0.0203
  -1.750   0.2779   0.02985   0.02414  -0.1132   0.8924   0.0225
  -1.500   0.3068   0.02790   0.02199  -0.1138   0.8910   0.0271
  -1.250   0.3205   0.02854   0.02240  -0.1104   0.8868   0.0321
  -1.000   0.3421   0.02622   0.01990  -0.1098   0.8831   0.0363
  -0.750   0.3730   0.02484   0.01826  -0.1102   0.8807   0.0490
  -0.500   0.4045   0.02340   0.01657  -0.1108   0.8789   0.0747
   0.000   0.4418   0.02081   0.01387  -0.1089   0.8707   0.1890
   0.250   0.4758   0.02056   0.01343  -0.1082   0.8681   0.1305
   0.500   0.5121   0.02072   0.01328  -0.1076   0.8661   0.0580
   0.750   0.5325   0.02027   0.01283  -0.1061   0.8614   0.0421
   1.000   0.5566   0.02017   0.01263  -0.1052   0.8562   0.0326
   1.250   0.5891   0.01931   0.01179  -0.1059   0.8533   0.0286
   1.500   0.6084   0.01931   0.01171  -0.1045   0.8466   0.0266
   1.750   0.6376   0.01911   0.01136  -0.1047   0.8420   0.0253
   2.000   0.6718   0.01884   0.01104  -0.1054   0.8392   0.0272
   2.250   0.6883   0.01812   0.01139  -0.1039   0.8308   0.5881
   2.500   0.7379   0.01700   0.01092  -0.1078   0.8278   1.0000
   2.750   0.7607   0.01716   0.01102  -0.1068   0.8207   1.0000
   3.000   0.7898   0.01695   0.01077  -0.1065   0.8153   1.0000
   3.250   0.8216   0.01241   0.00589  -0.0990   0.7662   1.0000
   3.500   0.8425   0.01042   0.00381  -0.0951   0.6772   1.0000
   3.750   0.8566   0.01048   0.00276  -0.0907   0.4907   1.0000
   4.000   0.8519   0.01346   0.00406  -0.0862   0.1915   1.0000
   4.250   0.8601   0.01553   0.00519  -0.0834   0.0138   1.0000
   4.500   0.8832   0.01608   0.00600  -0.0822   0.0133   1.0000
   4.750   0.9051   0.01681   0.00701  -0.0807   0.0146   1.0000
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