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

USA 27 AIRFOIL (usa27-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: USA 27 AIRFOIL (usa27-il)
Reynolds number: 1,000,000
Max Cl/Cd: 117.37 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa27-il-1000000.txt
Download as CSV file: xf-usa27-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 27 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2500   0.09472   0.09309  -0.0565   0.9926   0.0183
  -9.250  -0.2414   0.08961   0.08798  -0.0614   0.9900   0.0184
  -9.000  -0.2303   0.08459   0.08296  -0.0665   0.9877   0.0184
  -8.750  -0.2233   0.07878   0.07715  -0.0717   0.9857   0.0187
  -8.500  -0.2036   0.07581   0.07418  -0.0756   0.9844   0.0189
  -8.250  -0.1879   0.07271   0.07108  -0.0794   0.9797   0.0191
  -8.000  -0.1710   0.06892   0.06728  -0.0852   0.9742   0.0194
  -7.750  -0.1555   0.06516   0.06349  -0.0901   0.9675   0.0198
  -7.500  -0.1403   0.06122   0.05951  -0.0947   0.9595   0.0204
  -7.250  -0.1238   0.05238   0.05052  -0.1045   0.9465   0.0224
  -7.000  -0.1272   0.03779   0.03551  -0.1116   0.9326   0.0228
  -6.750  -0.1116   0.03560   0.03325  -0.1116   0.9230   0.0230
  -6.500  -0.0931   0.03375   0.03129  -0.1117   0.9139   0.0232
  -6.250  -0.0752   0.03192   0.02935  -0.1115   0.9040   0.0235
  -6.000  -0.0565   0.02994   0.02723  -0.1113   0.8951   0.0239
  -5.750  -0.0377   0.02779   0.02488  -0.1109   0.8855   0.0245
  -5.500  -0.0193   0.02543   0.02227  -0.1100   0.8757   0.0258
  -5.250   0.0038   0.02462   0.02091  -0.1084   0.8663   0.0274
  -5.000   0.0131   0.01993   0.01576  -0.1063   0.8563   0.0280
  -4.750   0.0352   0.01857   0.01434  -0.1056   0.8468   0.0284
  -4.500   0.0581   0.01760   0.01327  -0.1049   0.8363   0.0288
  -4.250   0.0809   0.01675   0.01229  -0.1041   0.8244   0.0293
  -4.000   0.1041   0.01597   0.01136  -0.1032   0.8118   0.0301
  -3.750   0.1276   0.01528   0.01051  -0.1022   0.7983   0.0311
  -3.500   0.1536   0.01566   0.01071  -0.1014   0.7822   0.0330
  -3.250   0.1729   0.01374   0.00844  -0.0998   0.7636   0.0346
  -3.000   0.1957   0.01305   0.00768  -0.0988   0.7384   0.0355
  -2.750   0.2172   0.01267   0.00712  -0.0974   0.7040   0.0365
  -2.500   0.2389   0.01244   0.00669  -0.0960   0.6728   0.0380
  -2.250   0.2625   0.01255   0.00663  -0.0949   0.6514   0.0399
  -1.750   0.3124   0.01100   0.00474  -0.0930   0.6248   0.0326
  -1.500   0.3371   0.01033   0.00400  -0.0922   0.6146   0.0313
  -1.250   0.3618   0.01006   0.00366  -0.0914   0.6045   0.0314
  -1.000   0.3871   0.00973   0.00329  -0.0908   0.5947   0.0311
  -0.750   0.4117   0.00949   0.00300  -0.0900   0.5840   0.0310
  -0.500   0.4360   0.00933   0.00277  -0.0892   0.5707   0.0313
  -0.250   0.4605   0.00921   0.00259  -0.0884   0.5559   0.0317
   0.000   0.4851   0.00912   0.00244  -0.0877   0.5407   0.0321
   0.250   0.5099   0.00906   0.00232  -0.0870   0.5247   0.0324
   0.500   0.5344   0.00903   0.00222  -0.0863   0.5064   0.0328
   0.750   0.5586   0.00905   0.00215  -0.0855   0.4859   0.0330
   1.000   0.5826   0.00911   0.00212  -0.0847   0.4661   0.0333
   1.500   0.6305   0.00918   0.00201  -0.0830   0.4329   0.0357
   1.750   0.6550   0.00925   0.00201  -0.0824   0.4189   0.0373
   2.000   0.6801   0.00931   0.00202  -0.0818   0.4077   0.0394
   2.250   0.7052   0.00939   0.00206  -0.0813   0.3984   0.0448
   2.500   0.7275   0.00910   0.00218  -0.0803   0.3900   0.2373
   2.750   0.7514   0.00902   0.00233  -0.0797   0.3826   0.3403
   3.250   0.8363   0.00811   0.00291  -0.0865   0.3674   0.9881
   3.500   0.8902   0.00834   0.00308  -0.0925   0.3597   0.9949
   3.750   0.9539   0.00853   0.00320  -0.1007   0.3516   0.9999
   4.000   0.9779   0.00864   0.00329  -0.1000   0.3466   1.0000
   4.250   0.9993   0.00878   0.00339  -0.0988   0.3412   1.0000
   4.500   1.0212   0.00889   0.00349  -0.0976   0.3359   1.0000
   4.750   1.0427   0.00902   0.00360  -0.0964   0.3292   1.0000
   5.000   1.0632   0.00918   0.00372  -0.0950   0.3220   1.0000
   5.250   1.0845   0.00930   0.00383  -0.0937   0.3149   1.0000
   5.500   1.1042   0.00948   0.00397  -0.0921   0.3079   1.0000
   5.750   1.1256   0.00959   0.00409  -0.0909   0.3026   1.0000
   6.000   1.1450   0.00977   0.00423  -0.0893   0.2945   1.0000
   6.250   1.1649   0.00993   0.00437  -0.0878   0.2847   1.0000
   6.500   1.1832   0.01015   0.00453  -0.0859   0.2704   1.0000
   6.750   1.2006   0.01042   0.00472  -0.0840   0.2536   1.0000
   7.000   1.2151   0.01082   0.00498  -0.0815   0.2275   1.0000
   7.250   1.2234   0.01152   0.00542  -0.0779   0.1864   1.0000
   7.500   1.2344   0.01211   0.00586  -0.0749   0.1637   1.0000
   7.750   1.2468   0.01255   0.00622  -0.0720   0.1484   1.0000
   8.000   1.2566   0.01311   0.00663  -0.0687   0.1246   1.0000
   8.250   1.2485   0.01468   0.00774  -0.0625   0.0556   1.0000
   8.500   1.2519   0.01581   0.00871  -0.0585   0.0192   1.0000
   8.750   1.2675   0.01631   0.00924  -0.0565   0.0169   1.0000
   9.000   1.2838   0.01679   0.00976  -0.0548   0.0158   1.0000
   9.250   1.2998   0.01728   0.01030  -0.0530   0.0150   1.0000
   9.500   1.3150   0.01784   0.01091  -0.0512   0.0142   1.0000
   9.750   1.3285   0.01850   0.01161  -0.0492   0.0135   1.0000
  10.000   1.3389   0.01937   0.01255  -0.0468   0.0127   1.0000
  10.250   1.3516   0.02011   0.01337  -0.0449   0.0124   1.0000
  10.500   1.3650   0.02083   0.01413  -0.0431   0.0121   1.0000
  10.750   1.3771   0.02163   0.01499  -0.0413   0.0117   1.0000
  11.000   1.3884   0.02253   0.01594  -0.0394   0.0113   1.0000
  11.250   1.3988   0.02349   0.01697  -0.0376   0.0109   1.0000
  11.500   1.4080   0.02457   0.01810  -0.0358   0.0106   1.0000
  11.750   1.4148   0.02587   0.01947  -0.0338   0.0102   1.0000
  12.000   1.4157   0.02765   0.02134  -0.0315   0.0099   1.0000
  12.250   1.4080   0.03023   0.02406  -0.0288   0.0096   1.0000
  12.500   1.4145   0.03178   0.02568  -0.0274   0.0095   1.0000
  12.750   1.4202   0.03347   0.02744  -0.0262   0.0094   1.0000
  13.000   1.4238   0.03543   0.02949  -0.0251   0.0092   1.0000
  13.250   1.4255   0.03769   0.03183  -0.0241   0.0090   1.0000
  13.500   1.4255   0.04023   0.03446  -0.0233   0.0089   1.0000
  13.750   1.4242   0.04308   0.03740  -0.0228   0.0087   1.0000
  14.000   1.4219   0.04619   0.04061  -0.0225   0.0086   1.0000
  14.250   1.4182   0.04958   0.04410  -0.0225   0.0085   1.0000
  14.500   1.4138   0.05316   0.04777  -0.0226   0.0083   1.0000
  14.750   1.4081   0.05697   0.05167  -0.0228   0.0082   1.0000
  15.000   1.4014   0.06098   0.05579  -0.0233   0.0081   1.0000
  15.250   1.3937   0.06517   0.06007  -0.0238   0.0081   1.0000
  15.500   1.3859   0.06951   0.06451  -0.0245   0.0080   1.0000
  15.750   1.3777   0.07397   0.06905  -0.0253   0.0079   1.0000
  16.000   1.3690   0.07852   0.07369  -0.0262   0.0078   1.0000
  16.250   1.3604   0.08315   0.07840  -0.0272   0.0077   1.0000
  16.500   1.3520   0.08781   0.08314  -0.0283   0.0076   1.0000
  16.750   1.3434   0.09249   0.08789  -0.0294   0.0076   1.0000
  17.000   1.3343   0.09717   0.09264  -0.0305   0.0075   1.0000
<< Back to USA 27 AIRFOIL (usa27-il)

Polar data table (+)

Polar graphs


<< Back to USA 27 AIRFOIL (usa27-il)