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

RG 14 9% AIRFOIL (rg149-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: RG 14 9% AIRFOIL (rg149-il)
Reynolds number: 200,000
Max Cl/Cd: 63.55 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rg149-il-200000.txt
Download as CSV file: xf-rg149-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 14 9% 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5215   0.08892   0.08539  -0.0307   1.0000   0.0568
  -9.000  -0.5406   0.08265   0.07916  -0.0382   1.0000   0.0571
  -8.750  -0.5585   0.07867   0.07512  -0.0398   1.0000   0.0572
  -8.500  -0.5728   0.07483   0.07116  -0.0406   1.0000   0.0573
  -8.250  -0.5719   0.06807   0.06455  -0.0401   1.0000   0.0587
  -8.000  -0.5629   0.06582   0.06236  -0.0382   1.0000   0.0599
  -7.750  -0.5601   0.06306   0.05958  -0.0372   1.0000   0.0613
  -7.500  -0.5603   0.05967   0.05612  -0.0368   1.0000   0.0629
  -7.250  -0.5611   0.05591   0.05224  -0.0364   1.0000   0.0653
  -7.000  -0.5788   0.05096   0.04657  -0.0355   1.0000   0.0711
  -6.750  -0.5667   0.04710   0.04287  -0.0344   1.0000   0.0727
  -6.500  -0.5704   0.03517   0.02998  -0.0308   1.0000   0.0424
  -6.250  -0.5602   0.03162   0.02600  -0.0286   1.0000   0.0419
  -6.000  -0.5474   0.02816   0.02209  -0.0264   1.0000   0.0411
  -5.750  -0.5316   0.02495   0.01838  -0.0243   1.0000   0.0403
  -5.500  -0.5129   0.02272   0.01577  -0.0226   1.0000   0.0407
  -5.250  -0.4928   0.02130   0.01407  -0.0211   1.0000   0.0425
  -5.000  -0.4719   0.02031   0.01282  -0.0196   1.0000   0.0446
  -4.750  -0.4505   0.01923   0.01153  -0.0182   1.0000   0.0454
  -4.500  -0.4292   0.01719   0.00938  -0.0171   1.0000   0.0476
  -4.250  -0.4081   0.01644   0.00865  -0.0160   1.0000   0.0510
  -4.000  -0.3866   0.01584   0.00798  -0.0148   1.0000   0.0551
  -3.750  -0.3653   0.01508   0.00717  -0.0136   1.0000   0.0583
  -3.500  -0.3449   0.01432   0.00648  -0.0125   1.0000   0.0648
  -3.250  -0.3242   0.01375   0.00593  -0.0114   1.0000   0.0767
  -3.000  -0.2986   0.01294   0.00536  -0.0114   0.9988   0.1196
  -2.750  -0.2595   0.01213   0.00512  -0.0144   0.9937   0.2356
  -2.500  -0.2233   0.01126   0.00501  -0.0169   0.9874   0.4089
  -2.250  -0.1936   0.01021   0.00520  -0.0173   0.9815   0.6786
  -2.000  -0.1647   0.01010   0.00552  -0.0163   0.9735   0.8338
  -1.750  -0.1316   0.01017   0.00560  -0.0164   0.9660   0.8916
  -1.500  -0.0899   0.01027   0.00565  -0.0182   0.9602   0.9314
  -1.250  -0.0304   0.01045   0.00573  -0.0238   0.9590   0.9619
  -1.000   0.0370   0.01054   0.00569  -0.0314   0.9590   0.9781
  -0.750   0.1037   0.01048   0.00554  -0.0391   0.9583   0.9896
  -0.500   0.1705   0.01031   0.00529  -0.0470   0.9574   0.9995
  -0.250   0.2142   0.01008   0.00502  -0.0503   0.9477   1.0000
   0.000   0.2598   0.00983   0.00474  -0.0539   0.9388   1.0000
   0.250   0.3042   0.00956   0.00445  -0.0572   0.9285   1.0000
   0.500   0.3418   0.00932   0.00419  -0.0590   0.9131   1.0000
   0.750   0.3733   0.00913   0.00397  -0.0595   0.8925   1.0000
   1.000   0.4039   0.00896   0.00374  -0.0597   0.8707   1.0000
   1.250   0.4275   0.00888   0.00360  -0.0585   0.8446   1.0000
   1.500   0.4500   0.00885   0.00350  -0.0571   0.8187   1.0000
   1.750   0.4716   0.00885   0.00342  -0.0556   0.7923   1.0000
   2.000   0.4927   0.00890   0.00338  -0.0539   0.7658   1.0000
   2.250   0.5132   0.00898   0.00336  -0.0522   0.7391   1.0000
   2.500   0.5333   0.00908   0.00338  -0.0505   0.7122   1.0000
   2.750   0.5533   0.00921   0.00345  -0.0488   0.6851   1.0000
   3.000   0.5730   0.00937   0.00351  -0.0470   0.6567   1.0000
   3.250   0.5922   0.00956   0.00358  -0.0451   0.6257   1.0000
   3.500   0.6115   0.00978   0.00368  -0.0432   0.5942   1.0000
   3.750   0.6309   0.01000   0.00383  -0.0414   0.5618   1.0000
   4.000   0.6506   0.01026   0.00398  -0.0397   0.5293   1.0000
   4.250   0.6702   0.01055   0.00416  -0.0380   0.4950   1.0000
   4.500   0.6901   0.01086   0.00436  -0.0364   0.4582   1.0000
   4.750   0.7100   0.01123   0.00461  -0.0349   0.4188   1.0000
   5.000   0.7297   0.01166   0.00488  -0.0333   0.3733   1.0000
   5.250   0.7488   0.01220   0.00520  -0.0318   0.3205   1.0000
   5.500   0.7675   0.01285   0.00560  -0.0303   0.2642   1.0000
   5.750   0.7864   0.01359   0.00610  -0.0289   0.2158   1.0000
   6.000   0.8059   0.01433   0.00665  -0.0276   0.1799   1.0000
   6.250   0.8255   0.01512   0.00730  -0.0263   0.1535   1.0000
   6.500   0.8448   0.01598   0.00804  -0.0250   0.1331   1.0000
   6.750   0.8658   0.01666   0.00872  -0.0240   0.1157   1.0000
   7.000   0.8867   0.01737   0.00943  -0.0229   0.1012   1.0000
   7.250   0.9067   0.01824   0.01028  -0.0217   0.0896   1.0000
   7.500   0.9260   0.01923   0.01123  -0.0204   0.0794   1.0000
   7.750   0.9464   0.02003   0.01208  -0.0193   0.0705   1.0000
   8.000   0.9662   0.02106   0.01319  -0.0181   0.0629   1.0000
   8.250   0.9842   0.02245   0.01453  -0.0169   0.0558   1.0000
   8.500   1.0050   0.02312   0.01540  -0.0158   0.0499   1.0000
   8.750   1.0223   0.02493   0.01717  -0.0145   0.0445   1.0000
   9.000   1.0417   0.02566   0.01812  -0.0133   0.0397   1.0000
   9.250   1.0588   0.02712   0.01957  -0.0120   0.0358   1.0000
   9.500   1.0760   0.02945   0.02217  -0.0106   0.0332   1.0000
   9.750   1.0926   0.03106   0.02404  -0.0091   0.0308   1.0000
  10.000   1.1075   0.03226   0.02534  -0.0077   0.0285   1.0000
  10.250   1.1187   0.03535   0.02856  -0.0064   0.0266   1.0000
  10.500   1.1256   0.03803   0.03162  -0.0041   0.0258   1.0000
  10.750   1.1297   0.04057   0.03452  -0.0016   0.0251   1.0000
  11.000   1.1280   0.04347   0.03779   0.0012   0.0248   1.0000
  11.250   1.1191   0.04640   0.04103   0.0047   0.0245   1.0000
  11.500   1.1058   0.04956   0.04449   0.0078   0.0244   1.0000
  11.750   1.0899   0.05309   0.04829   0.0099   0.0244   1.0000
  12.000   1.0705   0.05712   0.05257   0.0109   0.0244   1.0000
  12.250   1.0505   0.06176   0.05744   0.0105   0.0245   1.0000
  12.500   1.0275   0.06731   0.06321   0.0085   0.0246   1.0000
  12.750   1.0048   0.07372   0.06981   0.0050   0.0249   1.0000
  13.000   0.9806   0.08129   0.07754   0.0001   0.0251   1.0000
  13.250   0.9561   0.08997   0.08635  -0.0059   0.0255   1.0000
<< Back to RG 14 9% AIRFOIL (rg149-il)

Polar data table (+)

Polar graphs


<< Back to RG 14 9% AIRFOIL (rg149-il)