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GRUMMAN K-2 AIRFOIL (k2-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GRUMMAN K-2 AIRFOIL (k2-il)
Reynolds number: 500,000
Max Cl/Cd: 44.01 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-k2-il-500000.txt
Download as CSV file: xf-k2-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GRUMMAN K-2 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.7712   0.06480   0.06177  -0.0456   1.0000   0.0304
  -9.250  -0.8104   0.06197   0.05878  -0.0414   1.0000   0.0305
  -9.000  -0.8484   0.05958   0.05618  -0.0365   1.0000   0.0306
  -8.750  -0.8678   0.05707   0.05341  -0.0331   1.0000   0.0308
  -8.500  -0.9220   0.03964   0.03516  -0.0281   1.0000   0.0217
  -8.250  -0.9178   0.03643   0.03175  -0.0263   1.0000   0.0212
  -8.000  -0.9132   0.03231   0.02724  -0.0244   1.0000   0.0205
  -7.750  -0.9032   0.02829   0.02273  -0.0228   1.0000   0.0200
  -7.500  -0.8863   0.02597   0.02012  -0.0217   1.0000   0.0201
  -7.250  -0.8668   0.02429   0.01822  -0.0208   1.0000   0.0202
  -7.000  -0.8459   0.02284   0.01660  -0.0201   1.0000   0.0204
  -6.750  -0.8240   0.02163   0.01524  -0.0194   1.0000   0.0207
  -6.500  -0.8014   0.02056   0.01404  -0.0187   1.0000   0.0210
  -6.250  -0.7780   0.01961   0.01298  -0.0182   1.0000   0.0213
  -6.000  -0.7541   0.01880   0.01207  -0.0177   1.0000   0.0217
  -5.750  -0.7296   0.01816   0.01134  -0.0173   1.0000   0.0220
  -5.500  -0.7054   0.01699   0.01017  -0.0171   1.0000   0.0227
  -5.250  -0.6803   0.01643   0.00964  -0.0170   1.0000   0.0235
  -5.000  -0.6546   0.01598   0.00917  -0.0170   1.0000   0.0245
  -4.750  -0.6284   0.01552   0.00868  -0.0170   1.0000   0.0256
  -4.500  -0.6010   0.01483   0.00797  -0.0173   1.0000   0.0267
  -4.250  -0.5732   0.01429   0.00747  -0.0178   1.0000   0.0281
  -4.000  -0.5453   0.01391   0.00707  -0.0182   1.0000   0.0298
  -3.750  -0.5158   0.01338   0.00654  -0.0189   1.0000   0.0319
  -3.500  -0.4873   0.01305   0.00625  -0.0195   1.0000   0.0346
  -3.250  -0.4576   0.01267   0.00589  -0.0202   1.0000   0.0383
  -3.000  -0.4289   0.01246   0.00568  -0.0207   1.0000   0.0424
  -2.750  -0.3982   0.01209   0.00536  -0.0217   1.0000   0.0492
  -2.500  -0.3664   0.01171   0.00509  -0.0229   1.0000   0.0634
  -2.250  -0.3247   0.01074   0.00472  -0.0269   1.0000   0.1800
  -2.000  -0.2508   0.00869   0.00419  -0.0389   1.0000   0.5257
  -1.750  -0.2093   0.00830   0.00437  -0.0420   1.0000   0.6669
  -1.500  -0.1819   0.00842   0.00457  -0.0419   1.0000   0.7013
  -1.250  -0.1564   0.00861   0.00479  -0.0413   1.0000   0.7205
  -1.000  -0.1304   0.00881   0.00500  -0.0408   1.0000   0.7354
  -0.750  -0.0930   0.00918   0.00541  -0.0424   0.9963   0.7528
  -0.500  -0.0516   0.00964   0.00589  -0.0445   0.9903   0.7678
  -0.250  -0.0023   0.00980   0.00609  -0.0484   0.9841   0.7716
   0.000   0.0354   0.00977   0.00608  -0.0499   0.9762   0.7748
   0.250   0.0797   0.00964   0.00596  -0.0527   0.9691   0.7784
   0.500   0.1243   0.00944   0.00575  -0.0558   0.9615   0.7849
   0.750   0.1529   0.00945   0.00582  -0.0551   0.9549   0.7877
   1.000   0.1841   0.00943   0.00585  -0.0551   0.9488   0.7903
   1.250   0.2224   0.00927   0.00574  -0.0568   0.9446   0.7925
   1.500   0.2503   0.00910   0.00561  -0.0563   0.9345   0.7952
   1.750   0.2831   0.00884   0.00538  -0.0569   0.9210   0.7981
   2.000   0.3266   0.00849   0.00501  -0.0598   0.8917   0.8010
   2.250   0.3739   0.00879   0.00457  -0.0630   0.7041   0.8025
   2.500   0.3830   0.00993   0.00485  -0.0589   0.5065   0.8041
   2.750   0.3999   0.01082   0.00511  -0.0567   0.3539   0.8057
   3.000   0.4213   0.01148   0.00533  -0.0555   0.2539   0.8075
   3.250   0.4448   0.01204   0.00554  -0.0548   0.1768   0.8094
   3.500   0.4704   0.01240   0.00572  -0.0544   0.1386   0.8116
   3.750   0.4978   0.01269   0.00589  -0.0544   0.1160   0.8140
   4.000   0.5272   0.01292   0.00605  -0.0549   0.0997   0.8166
   4.250   0.5580   0.01316   0.00620  -0.0557   0.0817   0.8193
   4.500   0.5809   0.01370   0.00648  -0.0547   0.0389   0.8207
   4.750   0.6042   0.01431   0.00701  -0.0536   0.0267   0.8220
   5.000   0.6289   0.01478   0.00749  -0.0529   0.0230   0.8234
   5.250   0.6543   0.01514   0.00788  -0.0522   0.0210   0.8251
   5.500   0.6790   0.01567   0.00840  -0.0516   0.0195   0.8269
   5.750   0.7048   0.01616   0.00895  -0.0511   0.0185   0.8289
   6.000   0.7315   0.01667   0.00948  -0.0509   0.0176   0.8309
   6.250   0.7588   0.01724   0.01005  -0.0510   0.0169   0.8331
   6.500   0.7861   0.01813   0.01092  -0.0512   0.0163   0.8355
   6.750   0.8084   0.01902   0.01188  -0.0500   0.0160   0.8368
   7.000   0.8320   0.01975   0.01270  -0.0491   0.0157   0.8380
   7.250   0.8557   0.02062   0.01366  -0.0482   0.0156   0.8393
   7.500   0.8793   0.02164   0.01479  -0.0473   0.0155   0.8408
   7.750   0.9027   0.02280   0.01609  -0.0464   0.0155   0.8424
   8.000   0.9260   0.02409   0.01753  -0.0456   0.0154   0.8443
   8.250   0.9493   0.02545   0.01910  -0.0448   0.0153   0.8462
   8.500   0.9722   0.02695   0.02082  -0.0441   0.0152   0.8480
   8.750   0.9942   0.02870   0.02282  -0.0433   0.0151   0.8498
   9.000   1.0138   0.03091   0.02533  -0.0423   0.0152   0.8515
   9.250   1.0269   0.03340   0.02814  -0.0401   0.0154   0.8528
   9.500   1.0373   0.03610   0.03115  -0.0377   0.0156   0.8543
   9.750   1.0451   0.03935   0.03465  -0.0353   0.0159   0.8559
  10.000   0.8819   0.07122   0.06885  -0.0210   0.0255   0.8588
  10.250   0.8511   0.07889   0.07669  -0.0226   0.0253   0.8609
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