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

FX 66-182 AIRFOIL (fx66182-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: FX 66-182 AIRFOIL (fx66182-il)
Reynolds number: 200,000
Max Cl/Cd: 61.81 at α=11°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66182-il-200000.txt
Download as CSV file: xf-fx66182-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-182 AIRFOIL                               
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.2280   0.10379   0.10006  -0.0912   0.8236   0.0628
 -12.000  -0.2198   0.09868   0.09480  -0.0922   0.7989   0.0639
 -11.750  -0.2004   0.09757   0.09350  -0.0904   0.7773   0.0654
 -11.500  -0.1907   0.09524   0.09104  -0.0906   0.7628   0.0671
 -11.250  -0.1879   0.09195   0.08768  -0.0919   0.7524   0.0699
 -11.000  -0.1395   0.08100   0.07681  -0.0881   0.7184   0.0787
 -10.750  -0.1321   0.07828   0.07404  -0.0882   0.7128   0.0814
 -10.500  -0.1395   0.07322   0.06895  -0.0904   0.7096   0.0841
 -10.000  -0.3659   0.03720   0.03217  -0.1071   0.7160   0.0407
  -9.750  -0.4241   0.04670   0.04118  -0.1084   0.7207   0.0434
  -9.500  -0.4477   0.04033   0.03389  -0.1042   0.7158   0.0342
  -9.250  -0.4409   0.03750   0.03071  -0.1028   0.7105   0.0343
  -9.000  -0.4286   0.03478   0.02781  -0.1020   0.7047   0.0350
  -8.750  -0.4106   0.03307   0.02600  -0.1014   0.6991   0.0361
  -8.500  -0.3942   0.03107   0.02370  -0.1004   0.6946   0.0366
  -8.250  -0.3755   0.02912   0.02147  -0.0995   0.6900   0.0371
  -8.000  -0.3544   0.02750   0.01964  -0.0987   0.6853   0.0381
  -7.750  -0.3320   0.02615   0.01804  -0.0980   0.6812   0.0398
  -7.250  -0.2858   0.02365   0.01529  -0.0970   0.6734   0.0446
  -7.000  -0.2627   0.02288   0.01449  -0.0965   0.6691   0.0474
  -6.750  -0.2390   0.02229   0.01377  -0.0960   0.6653   0.0504
  -6.500  -0.2160   0.02138   0.01277  -0.0953   0.6622   0.0528
  -6.250  -0.1956   0.02058   0.01204  -0.0945   0.6587   0.0560
  -6.000  -0.1744   0.01998   0.01144  -0.0936   0.6551   0.0586
  -5.750  -0.1528   0.01944   0.01082  -0.0928   0.6519   0.0614
  -5.500  -0.1327   0.01877   0.01011  -0.0919   0.6491   0.0652
  -5.250  -0.1109   0.01827   0.00954  -0.0913   0.6465   0.0724
  -5.000  -0.0908   0.01762   0.00896  -0.0905   0.6433   0.0861
  -4.750  -0.0817   0.01569   0.00801  -0.0890   0.6400   0.2596
  -4.500  -0.0676   0.01482   0.00816  -0.0871   0.6370   0.4863
  -4.250  -0.0408   0.01519   0.00847  -0.0866   0.6343   0.5357
  -4.000  -0.0131   0.01563   0.00877  -0.0863   0.6319   0.5612
  -3.750   0.0139   0.01621   0.00927  -0.0857   0.6295   0.5825
  -3.500   0.0396   0.01676   0.00983  -0.0849   0.6264   0.5994
  -3.250   0.0658   0.01721   0.01025  -0.0842   0.6232   0.6138
  -3.000   0.0922   0.01767   0.01065  -0.0835   0.6203   0.6280
  -2.750   0.1187   0.01816   0.01116  -0.0825   0.6180   0.6399
  -2.500   0.1453   0.01859   0.01155  -0.0816   0.6159   0.6512
  -2.250   0.1709   0.01890   0.01180  -0.0811   0.6137   0.6621
  -2.000   0.1958   0.01906   0.01199  -0.0804   0.6106   0.6683
  -1.750   0.2224   0.01907   0.01189  -0.0808   0.6076   0.6751
  -1.500   0.2486   0.01909   0.01189  -0.0805   0.6050   0.6787
  -1.250   0.2759   0.01913   0.01187  -0.0805   0.6028   0.6827
  -1.000   0.3041   0.01916   0.01178  -0.0810   0.6009   0.6872
  -0.750   0.3329   0.01924   0.01172  -0.0817   0.5991   0.6912
  -0.500   0.3565   0.01933   0.01188  -0.0812   0.5962   0.6942
  -0.250   0.3817   0.01943   0.01200  -0.0811   0.5932   0.6975
   0.000   0.4084   0.01950   0.01204  -0.0812   0.5906   0.7014
   0.250   0.4364   0.01958   0.01203  -0.0818   0.5883   0.7056
   0.500   0.4638   0.01963   0.01204  -0.0820   0.5864   0.7085
   0.750   0.4918   0.01973   0.01211  -0.0822   0.5846   0.7115
   1.000   0.5182   0.01996   0.01232  -0.0824   0.5826   0.7147
   1.250   0.5411   0.02022   0.01264  -0.0821   0.5795   0.7185
   1.500   0.5668   0.02042   0.01284  -0.0823   0.5765   0.7218
   1.750   0.5927   0.02053   0.01298  -0.0823   0.5741   0.7243
   2.000   0.6194   0.02067   0.01312  -0.0824   0.5720   0.7275
   2.250   0.6476   0.02081   0.01323  -0.0828   0.5702   0.7308
   2.500   0.6770   0.02097   0.01334  -0.0834   0.5685   0.7341
   2.750   0.6989   0.02141   0.01385  -0.0831   0.5654   0.7375
   3.000   0.7191   0.02180   0.01436  -0.0823   0.5621   0.7404
   3.250   0.7433   0.02208   0.01470  -0.0821   0.5594   0.7437
   3.500   0.7705   0.02224   0.01487  -0.0824   0.5571   0.7471
   3.750   0.7990   0.02237   0.01499  -0.0829   0.5551   0.7507
   4.000   0.8283   0.02252   0.01512  -0.0835   0.5534   0.7542
   4.250   0.8523   0.02294   0.01561  -0.0833   0.5514   0.7577
   4.500   0.8632   0.02385   0.01674  -0.0814   0.5471   0.7615
   4.750   0.8848   0.02432   0.01729  -0.0810   0.5440   0.7655
   5.000   0.9117   0.02452   0.01751  -0.0813   0.5417   0.7694
   5.250   0.9396   0.02462   0.01767  -0.0815   0.5398   0.7734
   5.500   0.9683   0.02479   0.01788  -0.0819   0.5383   0.7784
   5.750   0.9986   0.02506   0.01815  -0.0827   0.5369   0.7834
   6.000   0.9870   0.02705   0.02048  -0.0781   0.5305   0.7881
   6.250   1.0056   0.02764   0.02118  -0.0772   0.5275   0.7935
   6.500   1.0340   0.02778   0.02138  -0.0777   0.5255   0.7997
   6.750   1.0644   0.02773   0.02140  -0.0782   0.5239   0.8058
   7.000   1.0952   0.02781   0.02154  -0.0789   0.5225   0.8137
   7.250   1.0458   0.03167   0.02575  -0.0699   0.5133   0.8234
   7.500   1.0703   0.03184   0.02603  -0.0697   0.5110   0.8341
   7.750   0.9996   0.03663   0.03103  -0.0591   0.5005   0.8542
   8.000   1.0651   0.03455   0.02911  -0.0629   0.5018   0.8787
   8.250   1.1653   0.03115   0.02585  -0.0717   0.5034   0.9608
   8.500   1.0371   0.04027   0.03505  -0.0576   0.4886   1.0000
   8.750   1.0923   0.03816   0.03298  -0.0596   0.4882   1.0000
   9.000   1.1549   0.03545   0.03033  -0.0623   0.4865   1.0000
   9.250   0.9707   0.05368   0.04850  -0.0529   0.4602   1.0000
   9.500   1.0208   0.05106   0.04594  -0.0532   0.4605   1.0000
   9.750   1.0739   0.04767   0.04264  -0.0535   0.4601   1.0000
  10.000   1.1252   0.04418   0.03923  -0.0535   0.4583   1.0000
  10.250   1.0766   0.05166   0.04672  -0.0516   0.4447   1.0000
  10.500   1.1258   0.04801   0.04316  -0.0513   0.4415   1.0000
  10.750   1.0514   0.05876   0.05389  -0.0496   0.4236   1.0000
  11.000   1.4377   0.02326   0.01831  -0.0635   0.4237   1.0000
  11.250   1.4392   0.02394   0.01907  -0.0604   0.4130   1.0000
  11.500   1.4343   0.02511   0.02030  -0.0572   0.3991   1.0000
  11.750   1.4260   0.02685   0.02207  -0.0543   0.3833   1.0000
  12.000   1.4158   0.02905   0.02426  -0.0518   0.3635   1.0000
  12.250   1.4018   0.03185   0.02701  -0.0496   0.3393   1.0000
  12.500   1.3836   0.03520   0.03022  -0.0475   0.3107   1.0000
  12.750   1.3599   0.03925   0.03410  -0.0455   0.2784   1.0000
  13.000   1.3329   0.04381   0.03846  -0.0437   0.2493   1.0000
  13.250   1.3081   0.04849   0.04298  -0.0424   0.2221   1.0000
  13.500   1.2820   0.05358   0.04788  -0.0414   0.1929   1.0000
  13.750   1.2572   0.05882   0.05293  -0.0408   0.1628   1.0000
  14.000   1.2347   0.06407   0.05798  -0.0406   0.1332   1.0000
  14.250   1.2119   0.06955   0.06322  -0.0405   0.1009   1.0000
  14.500   1.1932   0.07477   0.06822  -0.0406   0.0746   1.0000
  14.750   1.1766   0.07990   0.07318  -0.0409   0.0563   1.0000
  15.000   1.1662   0.08442   0.07763  -0.0413   0.0471   1.0000
  15.250   1.1622   0.08821   0.08145  -0.0417   0.0417   1.0000
  15.500   1.1565   0.09227   0.08551  -0.0422   0.0385   1.0000
<< Back to FX 66-182 AIRFOIL (fx66182-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-182 AIRFOIL (fx66182-il)