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AG47ct -02f rot. (ag47ct02r-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG47ct -02f rot. (ag47ct02r-il)
Reynolds number: 200,000
Max Cl/Cd: 52.1 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag47ct02r-il-200000.txt
Download as CSV file: xf-ag47ct02r-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG47ct  -02f rot.                               
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.5277   0.13541   0.13203   0.0241   1.0000   0.0265
 -10.750  -0.5276   0.13221   0.12884   0.0226   1.0000   0.0266
  -8.750  -0.6291   0.11004   0.10662   0.0307   1.0000   0.0287
  -8.500  -0.6232   0.10657   0.10313   0.0303   1.0000   0.0299
  -8.250  -0.6187   0.10295   0.09953   0.0288   1.0000   0.0307
  -8.000  -0.6144   0.09927   0.09588   0.0269   1.0000   0.0316
  -7.750  -0.6104   0.09552   0.09216   0.0247   1.0000   0.0325
  -7.500  -0.6062   0.09166   0.08833   0.0219   1.0000   0.0335
  -7.250  -0.5993   0.08740   0.08410   0.0175   1.0000   0.0346
  -7.000  -0.5872   0.08249   0.07919   0.0106   1.0000   0.0360
  -6.750  -0.5626   0.07651   0.07312  -0.0025   1.0000   0.0377
  -6.500  -0.5401   0.07135   0.06777  -0.0099   1.0000   0.0382
  -6.250  -0.5273   0.06422   0.06052  -0.0145   1.0000   0.0387
  -6.000  -0.5189   0.05963   0.05602  -0.0140   1.0000   0.0400
  -5.750  -0.5033   0.05632   0.05269  -0.0147   1.0000   0.0416
  -5.500  -0.4832   0.05250   0.04878  -0.0168   1.0000   0.0438
  -5.250  -0.4582   0.04811   0.04417  -0.0199   1.0000   0.0475
  -5.000  -0.4290   0.04234   0.03782  -0.0240   1.0000   0.0522
  -4.750  -0.4093   0.03913   0.03467  -0.0243   1.0000   0.0544
  -4.500  -0.3852   0.03624   0.03160  -0.0250   1.0000   0.0586
  -4.250  -0.3578   0.03249   0.02737  -0.0262   1.0000   0.0667
  -4.000  -0.3295   0.03157   0.02607  -0.0260   1.0000   0.0768
  -3.750  -0.2963   0.02204   0.01566  -0.0253   1.0000   0.0332
  -3.500  -0.2678   0.01920   0.01233  -0.0248   1.0000   0.0324
  -3.250  -0.2399   0.01748   0.01028  -0.0243   1.0000   0.0348
  -3.000  -0.2118   0.01550   0.00800  -0.0236   1.0000   0.0346
  -2.750  -0.1842   0.01403   0.00636  -0.0229   1.0000   0.0362
  -2.500  -0.1573   0.01256   0.00479  -0.0223   1.0000   0.0406
  -2.250  -0.1304   0.01170   0.00395  -0.0219   1.0000   0.0522
  -2.000  -0.1035   0.01070   0.00307  -0.0216   1.0000   0.0804
  -1.750  -0.0776   0.00907   0.00246  -0.0218   1.0000   0.2955
  -1.500  -0.0640   0.00707   0.00240  -0.0185   1.0000   0.7954
  -1.250  -0.0248   0.00673   0.00212  -0.0199   1.0000   1.0000
  -1.000   0.0017   0.00672   0.00196  -0.0196   1.0000   1.0000
  -0.750   0.0283   0.00672   0.00184  -0.0193   1.0000   1.0000
  -0.500   0.0547   0.00673   0.00176  -0.0190   1.0000   1.0000
  -0.250   0.0811   0.00674   0.00170  -0.0187   1.0000   1.0000
   0.000   0.1074   0.00677   0.00169  -0.0184   1.0000   1.0000
   0.250   0.1335   0.00681   0.00171  -0.0181   1.0000   1.0000
   0.500   0.1596   0.00687   0.00177  -0.0179   1.0000   1.0000
   0.750   0.2025   0.00694   0.00184  -0.0213   0.9785   1.0000
   1.000   0.2508   0.00698   0.00186  -0.0253   0.9276   1.0000
   1.250   0.2856   0.00708   0.00183  -0.0260   0.8606   1.0000
   1.500   0.3100   0.00731   0.00186  -0.0244   0.8029   1.0000
   1.750   0.3339   0.00760   0.00192  -0.0230   0.7576   1.0000
   2.000   0.3585   0.00791   0.00204  -0.0219   0.7199   1.0000
   2.250   0.3839   0.00818   0.00218  -0.0211   0.6835   1.0000
   2.500   0.4097   0.00837   0.00229  -0.0204   0.6419   1.0000
   2.750   0.4353   0.00859   0.00239  -0.0196   0.5917   1.0000
   3.000   0.4605   0.00890   0.00248  -0.0188   0.5303   1.0000
   3.250   0.4856   0.00932   0.00261  -0.0182   0.4581   1.0000
   3.500   0.5106   0.00986   0.00282  -0.0177   0.3764   1.0000
   3.750   0.5359   0.01050   0.00311  -0.0174   0.2966   1.0000
   4.000   0.5613   0.01119   0.00354  -0.0171   0.2311   1.0000
   4.250   0.5868   0.01191   0.00400  -0.0169   0.1828   1.0000
   4.500   0.6126   0.01260   0.00455  -0.0166   0.1496   1.0000
   4.750   0.6383   0.01334   0.00517  -0.0164   0.1258   1.0000
   5.000   0.6642   0.01399   0.00584  -0.0161   0.1067   1.0000
   5.250   0.6898   0.01479   0.00663  -0.0157   0.0927   1.0000
   5.500   0.7153   0.01558   0.00743  -0.0153   0.0800   1.0000
   5.750   0.7405   0.01661   0.00851  -0.0149   0.0702   1.0000
   6.000   0.7649   0.01796   0.00983  -0.0144   0.0616   1.0000
   6.250   0.7910   0.01858   0.01065  -0.0140   0.0538   1.0000
   6.500   0.8148   0.02056   0.01273  -0.0134   0.0488   1.0000
   6.750   0.8402   0.02184   0.01427  -0.0128   0.0442   1.0000
   7.000   0.8639   0.02322   0.01570  -0.0124   0.0394   1.0000
   7.250   0.8861   0.02628   0.01922  -0.0117   0.0372   1.0000
   7.500   0.9083   0.02904   0.02249  -0.0109   0.0356   1.0000
   7.750   0.9273   0.03286   0.02692  -0.0101   0.0345   1.0000
   8.000   0.9453   0.03599   0.03055  -0.0096   0.0323   1.0000
   8.250   0.9629   0.03862   0.03344  -0.0093   0.0305   1.0000
   8.500   0.9595   0.04836   0.04417  -0.0097   0.0319   1.0000
   8.750   0.9470   0.05861   0.05503  -0.0122   0.0341   1.0000
   9.000   0.9349   0.06669   0.06334  -0.0158   0.0355   1.0000
   9.250   0.9204   0.07392   0.07072  -0.0205   0.0366   1.0000
   9.500   0.9064   0.08086   0.07768  -0.0263   0.0377   1.0000
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