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

TSAGI 8% AIRFOIL (tsagi8-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: TSAGI 8% AIRFOIL (tsagi8-il)
Reynolds number: 50,000
Max Cl/Cd: 32.82 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-tsagi8-il-50000.txt
Download as CSV file: xf-tsagi8-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: TSAGI 8% AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5402   0.09722   0.09061   0.0082   1.0000   0.2805
  -8.000  -0.5387   0.09398   0.08744   0.0086   1.0000   0.2967
  -7.750  -0.5482   0.09147   0.08498   0.0087   1.0000   0.3144
  -7.500  -0.5329   0.08746   0.08099   0.0109   1.0000   0.3397
  -7.250  -0.5315   0.08460   0.07820   0.0130   1.0000   0.3694
  -7.000  -0.5132   0.08094   0.07453   0.0161   1.0000   0.4077
  -6.750  -0.5029   0.07776   0.07139   0.0191   1.0000   0.4459
  -5.500  -0.5394   0.04882   0.04077  -0.0127   1.0000   0.1484
  -5.250  -0.5217   0.04456   0.03640  -0.0116   1.0000   0.1420
  -5.000  -0.5054   0.04073   0.03195  -0.0098   1.0000   0.1306
  -4.750  -0.4882   0.03715   0.02766  -0.0076   1.0000   0.1225
  -4.500  -0.4683   0.03438   0.02428  -0.0054   1.0000   0.1196
  -4.250  -0.4473   0.03181   0.02118  -0.0035   1.0000   0.1205
  -4.000  -0.4258   0.02939   0.01877  -0.0023   1.0000   0.1297
  -3.750  -0.4011   0.02717   0.01610  -0.0008   1.0000   0.1357
  -3.500  -0.3741   0.02519   0.01389   0.0002   1.0000   0.1479
  -3.250  -0.1450   0.01753   0.00887  -0.0295   1.0000   1.0000
  -3.000  -0.1269   0.01725   0.00821  -0.0283   1.0000   1.0000
  -2.750  -0.1089   0.01701   0.00762  -0.0271   1.0000   1.0000
  -2.500  -0.0911   0.01682   0.00718  -0.0257   1.0000   1.0000
  -2.250  -0.0737   0.01666   0.00681  -0.0242   1.0000   1.0000
  -2.000  -0.0569   0.01655   0.00652  -0.0227   1.0000   1.0000
  -1.750  -0.0409   0.01647   0.00629  -0.0209   1.0000   1.0000
  -1.500  -0.0260   0.01644   0.00610  -0.0190   1.0000   1.0000
  -1.250  -0.0125   0.01645   0.00602  -0.0169   1.0000   1.0000
  -1.000  -0.0006   0.01653   0.00602  -0.0146   1.0000   1.0000
  -0.750   0.0088   0.01667   0.00610  -0.0119   1.0000   1.0000
  -0.500   0.0156   0.01690   0.00627  -0.0089   1.0000   1.0000
  -0.250   0.0207   0.01721   0.00651  -0.0058   1.0000   1.0000
   0.000   0.0255   0.01758   0.00680  -0.0028   1.0000   1.0000
   0.250   0.0309   0.01800   0.00713  -0.0001   1.0000   1.0000
   0.500   0.0372   0.01846   0.00753   0.0023   1.0000   1.0000
   0.750   0.0445   0.01896   0.00797   0.0044   1.0000   1.0000
   1.000   0.0527   0.01950   0.00847   0.0063   1.0000   1.0000
   1.250   0.0617   0.02009   0.00902   0.0079   1.0000   1.0000
   1.500   0.0941   0.02088   0.00982   0.0048   0.9920   1.0000
   1.750   0.1580   0.02185   0.01087  -0.0041   0.9717   1.0000
   2.000   0.2175   0.02266   0.01180  -0.0117   0.9513   1.0000
   2.250   0.2752   0.02333   0.01267  -0.0186   0.9305   1.0000
   2.500   0.3337   0.02389   0.01344  -0.0252   0.9099   1.0000
   2.750   0.3920   0.02431   0.01413  -0.0315   0.8888   1.0000
   3.000   0.4468   0.02460   0.01477  -0.0366   0.8674   1.0000
   3.250   0.4953   0.02483   0.01530  -0.0401   0.8454   1.0000
   3.500   0.5307   0.02521   0.01595  -0.0410   0.8224   1.0000
   3.750   0.5709   0.02534   0.01645  -0.0420   0.8003   1.0000
   4.000   0.5970   0.02572   0.01709  -0.0407   0.7759   1.0000
   4.250   0.6211   0.02607   0.01772  -0.0388   0.7508   1.0000
   4.500   0.6457   0.02612   0.01807  -0.0361   0.7232   1.0000
   4.750   0.6668   0.02375   0.01583  -0.0271   0.6648   1.0000
   5.000   0.6776   0.02176   0.01371  -0.0179   0.5925   1.0000
   5.250   0.6791   0.02069   0.01242  -0.0090   0.4740   1.0000
   5.500   0.6676   0.02375   0.01312  -0.0014   0.1923   1.0000
   5.750   0.6797   0.02622   0.01509   0.0013   0.1428   1.0000
   6.000   0.6998   0.02844   0.01707   0.0034   0.1244   1.0000
   6.250   0.7257   0.03069   0.01926   0.0047   0.1112   1.0000
   6.500   0.7508   0.03281   0.02160   0.0060   0.1013   1.0000
   6.750   0.7781   0.03606   0.02495   0.0069   0.0975   1.0000
   7.000   0.8004   0.03902   0.02841   0.0085   0.0966   1.0000
   7.250   0.8192   0.04230   0.03216   0.0103   0.0964   1.0000
   7.500   0.8332   0.04560   0.03600   0.0123   0.0957   1.0000
   7.750   0.8430   0.04908   0.04001   0.0144   0.0949   1.0000
   8.000   0.8507   0.05299   0.04434   0.0164   0.0956   1.0000
   8.250   0.8648   0.05791   0.04935   0.0174   0.0978   1.0000
   8.500   0.8441   0.06181   0.05427   0.0203   0.1038   1.0000
   8.750   0.8403   0.06677   0.05946   0.0213   0.1083   1.0000
   9.000   0.8258   0.07159   0.06459   0.0219   0.1138   1.0000
   9.250   0.7943   0.07689   0.07004   0.0214   0.1173   1.0000
   9.500   0.7909   0.08238   0.07558   0.0212   0.1254   1.0000
   9.750   0.7356   0.09118   0.08428   0.0128   0.1330   1.0000
  10.000   0.6455   0.08878   0.08222   0.0170   0.1371   1.0000
  10.250   0.5902   0.10075   0.09396   0.0066   0.1574   1.0000
<< Back to TSAGI 8% AIRFOIL (tsagi8-il)

Polar data table (+)

Polar graphs


<< Back to TSAGI 8% AIRFOIL (tsagi8-il)