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

FX 61-163 AIRFOIL (fx61163-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: FX 61-163 AIRFOIL (fx61163-il)
Reynolds number: 100,000
Max Cl/Cd: 50.02 at α=8.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx61163-il-100000-n5.txt
Download as CSV file: xf-fx61163-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 61-163 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.4180   0.09408   0.08919  -0.0530   1.0001   0.0230
 -11.000  -0.4330   0.08602   0.08113  -0.0579   1.0001   0.0224
 -10.750  -0.4541   0.07865   0.07373  -0.0624   1.0001   0.0220
 -10.500  -0.4770   0.07257   0.06757  -0.0655   1.0001   0.0216
 -10.250  -0.5000   0.06762   0.06254  -0.0670   1.0001   0.0213
 -10.000  -0.5219   0.06365   0.05849  -0.0671   1.0001   0.0211
  -9.750  -0.5438   0.06046   0.05522  -0.0658   1.0001   0.0209
  -9.500  -0.5673   0.05808   0.05279  -0.0630   1.0001   0.0208
  -9.250  -0.5777   0.05449   0.04902  -0.0634   0.9914   0.0206
  -9.000  -0.5702   0.04954   0.04364  -0.0668   0.9639   0.0203
  -8.750  -0.5548   0.04505   0.03865  -0.0696   0.9411   0.0201
  -8.500  -0.5302   0.04105   0.03413  -0.0725   0.9225   0.0201
  -8.250  -0.5000   0.03763   0.03024  -0.0754   0.9060   0.0202
  -8.000  -0.4666   0.03475   0.02691  -0.0779   0.8900   0.0205
  -7.750  -0.4323   0.03232   0.02413  -0.0799   0.8738   0.0209
  -7.500  -0.3982   0.03030   0.02180  -0.0813   0.8578   0.0215
  -7.250  -0.3664   0.02865   0.01990  -0.0820   0.8422   0.0222
  -6.750  -0.3108   0.02629   0.01716  -0.0820   0.8137   0.0255
  -6.500  -0.2872   0.02529   0.01612  -0.0817   0.8014   0.0273
  -6.250  -0.2639   0.02439   0.01513  -0.0813   0.7906   0.0290
  -6.000  -0.2404   0.02353   0.01411  -0.0810   0.7811   0.0310
  -5.750  -0.2182   0.02270   0.01314  -0.0808   0.7712   0.0334
  -5.500  -0.1950   0.02186   0.01221  -0.0809   0.7630   0.0387
  -5.250  -0.1715   0.02104   0.01132  -0.0810   0.7552   0.0497
  -5.000  -0.1491   0.01988   0.01040  -0.0813   0.7484   0.0924
  -4.750  -0.1349   0.01716   0.00966  -0.0827   0.7412   0.4395
  -4.500  -0.1122   0.01835   0.01123  -0.0795   0.7354   0.5700
  -4.250  -0.0876   0.02002   0.01281  -0.0766   0.7287   0.6146
  -4.000  -0.0633   0.02123   0.01385  -0.0738   0.7235   0.6319
  -3.750  -0.0352   0.02131   0.01370  -0.0741   0.7180   0.6430
  -3.500  -0.0095   0.02147   0.01370  -0.0733   0.7123   0.6469
  -3.250   0.0193   0.02139   0.01338  -0.0739   0.7076   0.6538
  -3.000   0.0464   0.02137   0.01321  -0.0740   0.7024   0.6596
  -2.750   0.0724   0.02145   0.01316  -0.0735   0.6972   0.6642
  -2.250   0.1285   0.02132   0.01272  -0.0744   0.6879   0.6752
  -2.000   0.1546   0.02134   0.01265  -0.0742   0.6829   0.6795
  -1.750   0.1860   0.02116   0.01228  -0.0759   0.6785   0.6868
  -1.500   0.2116   0.02123   0.01227  -0.0753   0.6744   0.6896
  -1.250   0.2371   0.02127   0.01227  -0.0750   0.6693   0.6933
  -1.000   0.2662   0.02121   0.01210  -0.0758   0.6649   0.6984
  -0.750   0.2964   0.02114   0.01189  -0.0768   0.6613   0.7028
  -0.500   0.3210   0.02121   0.01197  -0.0763   0.6563   0.7057
  -0.250   0.3476   0.02123   0.01195  -0.0763   0.6515   0.7093
   0.000   0.3783   0.02118   0.01179  -0.0775   0.6475   0.7140
   0.250   0.4062   0.02119   0.01177  -0.0779   0.6428   0.7178
   0.500   0.4312   0.02127   0.01186  -0.0776   0.6377   0.7206
   0.750   0.4588   0.02129   0.01184  -0.0778   0.6334   0.7239
   1.000   0.4885   0.02132   0.01181  -0.0786   0.6292   0.7281
   1.250   0.5165   0.02138   0.01191  -0.0793   0.6239   0.7320
   1.500   0.5424   0.02145   0.01198  -0.0791   0.6194   0.7345
   1.750   0.5708   0.02150   0.01199  -0.0793   0.6159   0.7375
   2.000   0.5966   0.02164   0.01220  -0.0795   0.6102   0.7412
   2.250   0.6272   0.02170   0.01227  -0.0807   0.6052   0.7452
   2.500   0.6552   0.02173   0.01229  -0.0809   0.6013   0.7477
   2.750   0.6788   0.02190   0.01256  -0.0804   0.5953   0.7504
   3.000   0.7055   0.02199   0.01271  -0.0805   0.5899   0.7536
   3.250   0.7361   0.02202   0.01271  -0.0813   0.5858   0.7569
   3.500   0.7637   0.02223   0.01303  -0.0821   0.5792   0.7606
   3.750   0.7888   0.02232   0.01321  -0.0817   0.5741   0.7629
   4.000   0.8165   0.02238   0.01329  -0.0818   0.5700   0.7656
   4.250   0.8399   0.02263   0.01370  -0.0815   0.5630   0.7687
   4.500   0.8689   0.02270   0.01382  -0.0821   0.5579   0.7720
   4.750   0.8971   0.02288   0.01410  -0.0827   0.5518   0.7755
   5.000   0.9206   0.02300   0.01435  -0.0821   0.5451   0.7779
   5.250   0.9478   0.02301   0.01441  -0.0821   0.5398   0.7806
   5.500   0.9706   0.02327   0.01486  -0.0817   0.5318   0.7837
   5.750   1.0004   0.02328   0.01491  -0.0823   0.5264   0.7870
   6.000   1.0247   0.02357   0.01539  -0.0823   0.5177   0.7903
   6.250   1.0514   0.02347   0.01537  -0.0820   0.5112   0.7927
   6.500   1.0718   0.02372   0.01582  -0.0811   0.5012   0.7955
   6.750   1.0967   0.02380   0.01602  -0.0809   0.4928   0.7987
   7.000   1.1226   0.02387   0.01620  -0.0808   0.4835   0.8020
   7.250   1.1463   0.02409   0.01660  -0.0806   0.4725   0.8051
   7.500   1.1680   0.02421   0.01685  -0.0797   0.4624   0.8077
   7.750   1.1914   0.02427   0.01700  -0.0791   0.4525   0.8108
   8.000   1.2121   0.02452   0.01740  -0.0783   0.4398   0.8142
   8.250   1.2336   0.02477   0.01776  -0.0776   0.4258   0.8177
   8.500   1.2531   0.02505   0.01815  -0.0766   0.4108   0.8207
   8.750   1.2695   0.02539   0.01856  -0.0750   0.3951   0.8237
   9.000   1.2840   0.02588   0.01914  -0.0734   0.3783   0.8273
   9.250   1.2970   0.02651   0.01985  -0.0716   0.3608   0.8313
   9.500   1.3078   0.02720   0.02062  -0.0696   0.3452   0.8351
   9.750   1.3165   0.02802   0.02154  -0.0674   0.3293   0.8387
  10.000   1.3224   0.02912   0.02268  -0.0652   0.3099   0.8427
  10.250   1.3251   0.03054   0.02413  -0.0631   0.2880   0.8471
  10.500   1.3232   0.03231   0.02589  -0.0609   0.2653   0.8515
  10.750   1.3193   0.03443   0.02800  -0.0590   0.2424   0.8564
  11.000   1.3144   0.03698   0.03049  -0.0578   0.2211   0.8615
  11.250   1.3102   0.03958   0.03311  -0.0566   0.2014   0.8664
  11.500   1.3053   0.04236   0.03590  -0.0557   0.1843   0.8716
  11.750   1.3009   0.04532   0.03886  -0.0551   0.1689   0.8772
  12.000   1.2955   0.04824   0.04184  -0.0543   0.1548   0.8830
  12.250   1.2900   0.05133   0.04497  -0.0538   0.1414   0.8901
  12.500   1.2836   0.05442   0.04811  -0.0531   0.1296   0.8984
  12.750   1.2766   0.05760   0.05135  -0.0525   0.1190   0.9089
  13.250   1.2574   0.06362   0.05754  -0.0504   0.1013   0.9896
  13.500   1.2539   0.06754   0.06150  -0.0514   0.0917   0.9999
  13.750   1.2474   0.07176   0.06571  -0.0523   0.0834   0.9999
  14.000   1.2438   0.07575   0.06977  -0.0533   0.0745   0.9999
  14.250   1.2390   0.07995   0.07405  -0.0544   0.0673   0.9999
  14.500   1.2320   0.08450   0.07860  -0.0557   0.0615   0.9999
  14.750   1.2282   0.08866   0.08284  -0.0568   0.0560   0.9999
  15.000   1.2219   0.09327   0.08750  -0.0582   0.0510   0.9999
  15.250   1.2170   0.09774   0.09203  -0.0597   0.0463   0.9999
  15.500   1.2128   0.10223   0.09663  -0.0613   0.0415   0.9999
  15.750   1.2058   0.10721   0.10163  -0.0632   0.0377   0.9999
  16.000   1.2023   0.11174   0.10629  -0.0650   0.0337   0.9999
<< Back to FX 61-163 AIRFOIL (fx61163-il)

Polar data table (+)

Polar graphs


<< Back to FX 61-163 AIRFOIL (fx61163-il)