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

GRUMMAN K-2 AIRFOIL (k2-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GRUMMAN K-2 AIRFOIL (k2-il)
Reynolds number: 100,000
Max Cl/Cd: 25.76 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-k2-il-100000-n5.txt
Download as CSV file: xf-k2-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.7526   0.07192   0.06549  -0.0473   1.0000   0.0308
  -9.750  -0.7710   0.06799   0.06146  -0.0471   1.0000   0.0307
  -9.500  -0.7905   0.06458   0.05794  -0.0456   1.0000   0.0306
  -9.250  -0.8104   0.06175   0.05500  -0.0429   1.0000   0.0305
  -9.000  -0.8269   0.05878   0.05188  -0.0403   1.0000   0.0304
  -8.750  -0.8382   0.05562   0.04849  -0.0380   1.0000   0.0304
  -8.500  -0.8454   0.05241   0.04501  -0.0357   1.0000   0.0304
  -8.250  -0.8476   0.04929   0.04158  -0.0336   1.0000   0.0304
  -8.000  -0.8458   0.04624   0.03819  -0.0316   1.0000   0.0304
  -7.750  -0.8400   0.04331   0.03488  -0.0298   1.0000   0.0305
  -7.500  -0.8305   0.04056   0.03173  -0.0282   1.0000   0.0307
  -7.250  -0.8172   0.03818   0.02906  -0.0269   1.0000   0.0310
  -7.000  -0.8010   0.03626   0.02698  -0.0258   1.0000   0.0314
  -6.750  -0.7835   0.03451   0.02504  -0.0248   1.0000   0.0319
  -6.500  -0.7647   0.03302   0.02342  -0.0238   1.0000   0.0327
  -6.250  -0.7452   0.03161   0.02182  -0.0229   1.0000   0.0340
  -6.000  -0.7245   0.03012   0.02004  -0.0218   1.0000   0.0355
  -5.750  -0.7034   0.02871   0.01847  -0.0209   1.0000   0.0366
  -5.500  -0.6822   0.02751   0.01726  -0.0200   1.0000   0.0376
  -5.250  -0.6606   0.02639   0.01609  -0.0190   1.0000   0.0388
  -5.000  -0.6387   0.02533   0.01494  -0.0180   1.0000   0.0403
  -4.750  -0.6164   0.02440   0.01390  -0.0170   1.0000   0.0425
  -4.500  -0.5947   0.02348   0.01309  -0.0164   1.0000   0.0450
  -4.250  -0.5720   0.02267   0.01224  -0.0157   1.0000   0.0481
  -4.000  -0.5490   0.02177   0.01134  -0.0152   1.0000   0.0508
  -3.750  -0.5251   0.02103   0.01063  -0.0149   1.0000   0.0554
  -3.500  -0.5002   0.02028   0.00989  -0.0149   1.0000   0.0610
  -3.250  -0.4745   0.01961   0.00920  -0.0149   1.0000   0.0687
  -3.000  -0.4481   0.01889   0.00857  -0.0151   1.0000   0.0815
  -2.750  -0.4208   0.01805   0.00799  -0.0155   1.0000   0.1120
  -2.500  -0.3906   0.01649   0.00744  -0.0175   1.0000   0.2779
  -2.250  -0.3648   0.01513   0.00757  -0.0177   1.0000   0.5419
  -2.000  -0.3591   0.01573   0.00863  -0.0110   1.0000   0.6420
  -1.750  -0.3404   0.01625   0.00918  -0.0082   1.0000   0.6952
  -1.500  -0.3226   0.01703   0.00995  -0.0050   1.0000   0.7380
  -1.000  -0.2937   0.01852   0.01138   0.0031   1.0000   0.7854
  -0.750  -0.2779   0.01877   0.01160   0.0063   1.0000   0.7935
  -0.500  -0.2542   0.01890   0.01165   0.0071   1.0000   0.8041
  -0.250  -0.2369   0.01908   0.01180   0.0098   1.0000   0.8137
   0.000  -0.2169   0.01919   0.01188   0.0116   1.0000   0.8235
   0.250  -0.1952   0.01925   0.01192   0.0128   1.0000   0.8321
   0.500  -0.1709   0.01923   0.01188   0.0132   1.0000   0.8381
   0.750  -0.1490   0.01920   0.01184   0.0143   1.0000   0.8429
   1.000  -0.1225   0.01919   0.01182   0.0140   1.0000   0.8479
   1.250  -0.0878   0.01926   0.01189   0.0120   0.9974   0.8519
   1.500  -0.0466   0.01937   0.01202   0.0092   0.9890   0.8543
   1.750  -0.0027   0.01944   0.01213   0.0058   0.9787   0.8567
   2.000   0.0441   0.01938   0.01212   0.0020   0.9649   0.8588
   2.250   0.0909   0.01915   0.01198  -0.0016   0.9485   0.8609
   2.500   0.1353   0.01871   0.01162  -0.0043   0.9262   0.8630
   2.750   0.1788   0.01815   0.01115  -0.0067   0.8999   0.8650
   3.000   0.2071   0.01771   0.01082  -0.0062   0.8703   0.8677
   3.250   0.2372   0.01725   0.01046  -0.0057   0.8342   0.8700
   3.500   0.3306   0.01670   0.00939  -0.0160   0.6390   0.8682
   3.750   0.3591   0.01760   0.00940  -0.0157   0.4506   0.8700
   4.000   0.3806   0.01833   0.00959  -0.0146   0.3362   0.8727
   4.250   0.4041   0.01892   0.00984  -0.0140   0.2609   0.8758
   4.500   0.4289   0.01951   0.01015  -0.0139   0.2055   0.8788
   4.750   0.4546   0.02007   0.01050  -0.0139   0.1649   0.8813
   5.000   0.4770   0.02043   0.01078  -0.0130   0.1382   0.8837
   5.250   0.5000   0.02088   0.01109  -0.0122   0.1125   0.8865
   5.500   0.5241   0.02139   0.01150  -0.0118   0.0866   0.8891
   5.750   0.5492   0.02202   0.01206  -0.0116   0.0644   0.8915
   6.000   0.5747   0.02272   0.01270  -0.0115   0.0505   0.8942
   6.250   0.5971   0.02342   0.01338  -0.0106   0.0436   0.8967
   6.500   0.6189   0.02415   0.01417  -0.0095   0.0390   0.8992
   6.750   0.6418   0.02496   0.01504  -0.0089   0.0352   0.9015
   7.000   0.6655   0.02583   0.01595  -0.0085   0.0324   0.9038
   7.250   0.6896   0.02698   0.01718  -0.0081   0.0308   0.9063
   7.500   0.7144   0.02823   0.01861  -0.0078   0.0294   0.9087
   7.750   0.7361   0.02941   0.01992  -0.0068   0.0283   0.9111
   8.000   0.7585   0.03073   0.02134  -0.0062   0.0275   0.9133
   8.250   0.7811   0.03240   0.02324  -0.0057   0.0266   0.9154
   8.500   0.8026   0.03429   0.02553  -0.0049   0.0255   0.9177
   8.750   0.8230   0.03601   0.02754  -0.0043   0.0243   0.9202
   9.000   0.8410   0.03748   0.02917  -0.0034   0.0235   0.9228
   9.250   0.8561   0.03940   0.03132  -0.0021   0.0228   0.9253
   9.500   0.8669   0.04232   0.03478  -0.0004   0.0226   0.9278
   9.750   0.8735   0.04571   0.03869   0.0014   0.0224   0.9303
  10.000   0.8752   0.04952   0.04300   0.0033   0.0223   0.9329
  10.250   0.8699   0.05335   0.04727   0.0056   0.0223   0.9360
  10.500   0.8588   0.05740   0.05170   0.0078   0.0224   0.9396
  10.750   0.8424   0.06168   0.05630   0.0094   0.0225   0.9432
  11.000   0.8224   0.06674   0.06164   0.0093   0.0226   0.9463
  11.250   0.8003   0.07273   0.06786   0.0067   0.0227   0.9493
  11.500   0.7807   0.08035   0.07565   0.0006   0.0229   0.9521
<< Back to GRUMMAN K-2 AIRFOIL (k2-il)

Polar data table (+)

Polar graphs


<< Back to GRUMMAN K-2 AIRFOIL (k2-il)