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=1,000,000 Ncrit=5


Details Polar file
Airfoil: TSAGI 8% AIRFOIL (tsagi8-il)
Reynolds number: 1,000,000
Max Cl/Cd: 69.67 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-tsagi8-il-1000000-n5.txt
Download as CSV file: xf-tsagi8-il-1000000-n5.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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5842   0.08361   0.08208  -0.0096   1.0000   0.0026
  -9.000  -0.5892   0.07873   0.07721  -0.0130   1.0000   0.0026
  -8.750  -0.6048   0.07255   0.07106  -0.0187   1.0000   0.0026
  -7.750  -0.7430   0.02224   0.01881  -0.0098   1.0000   0.0022
  -7.500  -0.7313   0.01980   0.01602  -0.0071   1.0000   0.0023
  -7.250  -0.7170   0.01792   0.01382  -0.0047   1.0000   0.0024
  -7.000  -0.6976   0.01630   0.01192  -0.0033   0.9992   0.0026
  -6.500  -0.6413   0.01397   0.00917  -0.0040   0.9900   0.0030
  -6.250  -0.6116   0.01317   0.00819  -0.0046   0.9848   0.0031
  -6.000  -0.5816   0.01208   0.00692  -0.0053   0.9786   0.0033
  -5.750  -0.5495   0.01133   0.00608  -0.0065   0.9709   0.0036
  -5.500  -0.5147   0.01073   0.00539  -0.0082   0.9602   0.0038
  -5.250  -0.4782   0.01041   0.00503  -0.0103   0.9434   0.0044
  -5.000  -0.4472   0.00999   0.00447  -0.0110   0.9167   0.0049
  -4.750  -0.4219   0.00968   0.00398  -0.0104   0.8870   0.0051
  -4.500  -0.3986   0.00928   0.00342  -0.0094   0.8600   0.0058
  -4.250  -0.3746   0.00903   0.00306  -0.0085   0.8368   0.0066
  -4.000  -0.3501   0.00886   0.00277  -0.0078   0.8150   0.0075
  -3.750  -0.3251   0.00870   0.00251  -0.0071   0.7956   0.0086
  -3.500  -0.3004   0.00847   0.00220  -0.0064   0.7771   0.0122
  -3.250  -0.2752   0.00831   0.00199  -0.0059   0.7600   0.0178
  -2.750  -0.2242   0.00806   0.00165  -0.0049   0.7279   0.0325
  -2.500  -0.1985   0.00795   0.00149  -0.0045   0.7126   0.0405
  -2.250  -0.1730   0.00779   0.00135  -0.0040   0.6985   0.0603
  -2.000  -0.1479   0.00760   0.00123  -0.0035   0.6852   0.0922
  -1.750  -0.1221   0.00749   0.00112  -0.0031   0.6719   0.1135
  -1.250  -0.0713   0.00716   0.00094  -0.0023   0.6463   0.1901
  -1.000  -0.0463   0.00697   0.00086  -0.0018   0.6342   0.2440
  -0.750  -0.0216   0.00675   0.00079  -0.0012   0.6223   0.3060
  -0.500  -0.0001   0.00626   0.00071  -0.0001   0.6111   0.4449
  -0.250   0.0170   0.00558   0.00063   0.0019   0.6006   0.6353
   0.000   0.0347   0.00513   0.00059   0.0041   0.5896   0.7608
   0.250   0.0565   0.00492   0.00059   0.0056   0.5785   0.8340
   0.500   0.0865   0.00480   0.00068   0.0053   0.5665   0.9092
   0.750   0.1356   0.00494   0.00083   0.0006   0.5519   0.9455
   1.000   0.1696   0.00507   0.00091  -0.0007   0.5386   0.9560
   1.250   0.2018   0.00521   0.00101  -0.0017   0.5247   0.9644
   1.500   0.2368   0.00537   0.00111  -0.0033   0.5089   0.9686
   1.750   0.2697   0.00551   0.00118  -0.0045   0.4896   0.9707
   2.000   0.2999   0.00563   0.00124  -0.0051   0.4706   0.9730
   2.250   0.3266   0.00575   0.00131  -0.0049   0.4524   0.9761
   2.500   0.3555   0.00590   0.00138  -0.0052   0.4296   0.9778
   2.750   0.3868   0.00610   0.00147  -0.0062   0.3962   0.9788
   3.000   0.4182   0.00626   0.00157  -0.0071   0.3748   0.9801
   3.250   0.4487   0.00644   0.00168  -0.0079   0.3476   0.9816
   3.500   0.4743   0.00710   0.00189  -0.0080   0.2413   0.9838
   3.750   0.4971   0.00773   0.00219  -0.0073   0.1559   0.9867
   4.000   0.5253   0.00838   0.00250  -0.0080   0.0751   0.9876
   4.250   0.5541   0.00887   0.00279  -0.0086   0.0306   0.9886
   4.500   0.5835   0.00919   0.00305  -0.0092   0.0146   0.9896
   4.750   0.6129   0.00942   0.00331  -0.0097   0.0106   0.9908
   5.000   0.6413   0.00976   0.00366  -0.0100   0.0074   0.9921
   5.250   0.6694   0.01002   0.00396  -0.0103   0.0066   0.9933
   5.500   0.6971   0.01033   0.00431  -0.0105   0.0059   0.9946
   5.750   0.7253   0.01075   0.00476  -0.0109   0.0050   0.9955
   6.000   0.7537   0.01121   0.00531  -0.0113   0.0045   0.9964
   6.250   0.7821   0.01156   0.00571  -0.0117   0.0040   0.9973
   6.500   0.8104   0.01197   0.00616  -0.0122   0.0035   0.9983
   6.750   0.8384   0.01242   0.00666  -0.0126   0.0033   0.9992
   7.000   0.8647   0.01314   0.00745  -0.0128   0.0029   1.0000
   7.250   0.8820   0.01397   0.00838  -0.0109   0.0028   1.0000
   7.500   0.9014   0.01451   0.00900  -0.0095   0.0027   1.0000
   7.750   0.9196   0.01520   0.00980  -0.0079   0.0026   1.0000
   8.000   0.9364   0.01604   0.01079  -0.0060   0.0024   1.0000
   8.250   0.9528   0.01694   0.01182  -0.0041   0.0023   1.0000
   8.500   0.9676   0.01807   0.01309  -0.0020   0.0022   1.0000
   8.750   0.9821   0.01924   0.01442   0.0002   0.0021   1.0000
   9.000   0.9974   0.02026   0.01557   0.0021   0.0020   1.0000
   9.250   1.0140   0.02100   0.01641   0.0037   0.0019   1.0000
   9.500   1.0255   0.02248   0.01808   0.0061   0.0019   1.0000
   9.750   1.0399   0.02340   0.01912   0.0079   0.0018   1.0000
  10.000   1.0551   0.02408   0.01987   0.0095   0.0017   1.0000
  10.250   1.0647   0.02548   0.02144   0.0119   0.0017   1.0000
  10.500   1.0674   0.02762   0.02383   0.0150   0.0016   1.0000
  10.750   1.0540   0.03137   0.02799   0.0198   0.0015   1.0000
  11.000   1.0386   0.03418   0.03108   0.0249   0.0015   1.0000
  11.500   1.0067   0.04037   0.03771   0.0309   0.0015   1.0000
  12.000   0.9758   0.04891   0.04664   0.0300   0.0016   1.0000
  12.250   0.9473   0.05639   0.05435   0.0261   0.0015   1.0000
<< Back to TSAGI 8% AIRFOIL (tsagi8-il)

Polar data table (+)

Polar graphs


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