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

WORTMANN FX 60-126/1 AIRFOIL (fx601261-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: WORTMANN FX 60-126/1 AIRFOIL (fx601261-il)
Reynolds number: 500,000
Max Cl/Cd: 84.22 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx601261-il-500000-n5.txt
Download as CSV file: xf-fx601261-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX 60-126/1 AIRFOIL                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.3413   0.10042   0.09813  -0.0408   1.0000   0.0119
 -10.000  -0.3403   0.09709   0.09483  -0.0416   1.0000   0.0120
  -8.750  -0.3469   0.06374   0.06158  -0.0647   0.9663   0.0127
  -8.250  -0.3255   0.02933   0.02541  -0.1216   0.9053   0.0064
  -8.000  -0.3018   0.02640   0.02210  -0.1247   0.8904   0.0063
  -7.750  -0.2798   0.02404   0.01939  -0.1261   0.8777   0.0063
  -7.500  -0.2569   0.02217   0.01719  -0.1268   0.8661   0.0062
  -7.250  -0.2333   0.02062   0.01536  -0.1270   0.8551   0.0063
  -7.000  -0.2089   0.01932   0.01381  -0.1271   0.8454   0.0064
  -6.750  -0.1835   0.01826   0.01252  -0.1271   0.8370   0.0065
  -6.500  -0.1577   0.01735   0.01144  -0.1271   0.8294   0.0066
  -6.250  -0.1312   0.01664   0.01054  -0.1271   0.8226   0.0067
  -6.000  -0.1045   0.01595   0.00973  -0.1271   0.8161   0.0067
  -5.750  -0.0795   0.01471   0.00844  -0.1270   0.8091   0.0070
  -5.500  -0.0531   0.01401   0.00770  -0.1271   0.8020   0.0074
  -5.250  -0.0262   0.01345   0.00709  -0.1271   0.7939   0.0078
  -5.000   0.0012   0.01290   0.00647  -0.1271   0.7866   0.0078
  -4.750   0.0288   0.01236   0.00585  -0.1272   0.7789   0.0077
  -4.500   0.0568   0.01188   0.00531  -0.1274   0.7720   0.0076
  -4.250   0.0852   0.01144   0.00482  -0.1276   0.7646   0.0075
  -4.000   0.1135   0.01106   0.00436  -0.1277   0.7572   0.0075
  -3.750   0.1420   0.01072   0.00396  -0.1279   0.7483   0.0075
  -3.500   0.1704   0.01044   0.00360  -0.1280   0.7399   0.0075
  -3.250   0.1989   0.01021   0.00330  -0.1280   0.7310   0.0075
  -3.000   0.2274   0.01002   0.00302  -0.1281   0.7229   0.0075
  -2.750   0.2558   0.00987   0.00279  -0.1281   0.7135   0.0076
  -2.500   0.2842   0.00973   0.00258  -0.1281   0.7036   0.0077
  -2.250   0.3124   0.00964   0.00240  -0.1280   0.6932   0.0078
  -2.000   0.3406   0.00956   0.00225  -0.1280   0.6825   0.0080
  -1.750   0.3689   0.00950   0.00213  -0.1279   0.6726   0.0083
  -1.500   0.3970   0.00946   0.00203  -0.1278   0.6629   0.0090
  -1.250   0.4255   0.00934   0.00192  -0.1279   0.6524   0.0409
  -1.000   0.4543   0.00904   0.00179  -0.1282   0.6415   0.1161
  -0.750   0.4836   0.00841   0.00168  -0.1291   0.6290   0.2983
  -0.500   0.5115   0.00827   0.00174  -0.1292   0.6148   0.3791
  -0.250   0.5388   0.00832   0.00178  -0.1290   0.5975   0.4063
   0.000   0.5660   0.00838   0.00185  -0.1288   0.5793   0.4404
   0.250   0.5930   0.00849   0.00190  -0.1285   0.5604   0.4542
   0.500   0.6201   0.00859   0.00196  -0.1283   0.5449   0.4649
   0.750   0.6475   0.00870   0.00201  -0.1281   0.5311   0.4727
   1.000   0.6747   0.00882   0.00208  -0.1278   0.5157   0.4799
   1.250   0.7013   0.00898   0.00216  -0.1275   0.4959   0.4883
   1.500   0.7273   0.00918   0.00226  -0.1271   0.4706   0.4966
   1.750   0.7532   0.00940   0.00237  -0.1267   0.4443   0.5045
   2.000   0.7790   0.00963   0.00250  -0.1263   0.4190   0.5120
   2.250   0.8051   0.00985   0.00264  -0.1259   0.3982   0.5209
   2.500   0.8313   0.01006   0.00279  -0.1256   0.3764   0.5295
   2.750   0.8568   0.01034   0.00296  -0.1252   0.3525   0.5379
   3.000   0.8824   0.01060   0.00315  -0.1248   0.3291   0.5458
   3.250   0.9080   0.01088   0.00335  -0.1244   0.3072   0.5540
   3.500   0.9336   0.01114   0.00355  -0.1240   0.2891   0.5624
   3.750   0.9593   0.01139   0.00375  -0.1236   0.2726   0.5711
   4.000   0.9844   0.01169   0.00400  -0.1232   0.2502   0.5811
   4.250   1.0091   0.01202   0.00427  -0.1227   0.2312   0.5922
   4.500   1.0341   0.01232   0.00453  -0.1222   0.2139   0.6045
   5.000   1.0825   0.01306   0.00515  -0.1211   0.1749   0.6307
   5.250   1.1072   0.01335   0.00546  -0.1206   0.1610   0.6450
   5.500   1.1310   0.01372   0.00581  -0.1199   0.1430   0.6619
   5.750   1.1546   0.01409   0.00617  -0.1193   0.1313   0.6818
   6.000   1.1777   0.01448   0.00657  -0.1186   0.1164   0.7061
   6.250   1.1993   0.01495   0.00702  -0.1177   0.0964   0.7387
   6.500   1.2217   0.01524   0.00740  -0.1168   0.0887   0.7870
   6.750   1.2354   0.01524   0.00767  -0.1138   0.0818   1.0000
   7.000   1.2574   0.01574   0.00811  -0.1129   0.0666   1.0000
   7.250   1.2787   0.01627   0.00856  -0.1120   0.0600   1.0000
   7.500   1.3010   0.01668   0.00898  -0.1112   0.0560   1.0000
   7.750   1.3229   0.01711   0.00941  -0.1103   0.0520   1.0000
   8.000   1.3433   0.01763   0.00992  -0.1092   0.0475   1.0000
   8.250   1.3650   0.01802   0.01034  -0.1083   0.0389   1.0000
   8.500   1.3805   0.01883   0.01105  -0.1065   0.0312   1.0000
   8.750   1.3970   0.01945   0.01167  -0.1048   0.0282   1.0000
   9.000   1.4129   0.02003   0.01227  -0.1030   0.0264   1.0000
   9.250   1.4289   0.02057   0.01284  -0.1012   0.0251   1.0000
   9.500   1.4455   0.02107   0.01340  -0.0996   0.0242   1.0000
   9.750   1.4612   0.02163   0.01401  -0.0978   0.0227   1.0000
  10.000   1.4752   0.02230   0.01472  -0.0960   0.0191   1.0000
  10.500   1.4820   0.02511   0.01749  -0.0900   0.0030   1.0000
  10.750   1.4923   0.02614   0.01858  -0.0881   0.0028   1.0000
  11.000   1.5021   0.02725   0.01978  -0.0863   0.0026   1.0000
  11.250   1.5118   0.02841   0.02102  -0.0846   0.0025   1.0000
  11.500   1.5206   0.02970   0.02240  -0.0830   0.0024   1.0000
  11.750   1.5292   0.03105   0.02384  -0.0815   0.0024   1.0000
  12.000   1.5370   0.03254   0.02542  -0.0801   0.0023   1.0000
  12.250   1.5441   0.03413   0.02711  -0.0788   0.0023   1.0000
  12.500   1.5499   0.03592   0.02901  -0.0776   0.0022   1.0000
  12.750   1.5550   0.03784   0.03104  -0.0764   0.0022   1.0000
  13.000   1.5587   0.03997   0.03328  -0.0754   0.0021   1.0000
  13.250   1.5616   0.04227   0.03570  -0.0746   0.0020   1.0000
  13.500   1.5630   0.04482   0.03838  -0.0738   0.0019   1.0000
  13.750   1.5636   0.04755   0.04123  -0.0733   0.0019   1.0000
  14.000   1.5632   0.05051   0.04432  -0.0730   0.0019   1.0000
  14.250   1.5617   0.05373   0.04767  -0.0728   0.0018   1.0000
  14.500   1.5592   0.05722   0.05129  -0.0730   0.0018   1.0000
  14.750   1.5552   0.06105   0.05527  -0.0733   0.0018   1.0000
  15.000   1.5500   0.06523   0.05961  -0.0740   0.0017   1.0000
  15.250   1.5440   0.06968   0.06421  -0.0750   0.0017   1.0000
  15.500   1.5355   0.07471   0.06939  -0.0763   0.0017   1.0000
  15.750   1.5257   0.08016   0.07499  -0.0780   0.0017   1.0000
  16.000   1.5180   0.08543   0.08040  -0.0799   0.0017   1.0000
  16.250   1.5094   0.09099   0.08611  -0.0820   0.0017   1.0000
  16.500   1.4996   0.09694   0.09220  -0.0844   0.0016   1.0000
  16.750   1.4891   0.10318   0.09859  -0.0871   0.0016   1.0000
  17.000   1.4770   0.10987   0.10543  -0.0902   0.0016   1.0000
  17.250   1.4649   0.11670   0.11241  -0.0935   0.0016   1.0000
  17.500   1.4520   0.12387   0.11973  -0.0972   0.0016   1.0000
  17.750   1.4387   0.13124   0.12723  -0.1011   0.0016   1.0000
  18.000   1.4252   0.13877   0.13491  -0.1053   0.0016   1.0000
  18.250   1.4119   0.14640   0.14268  -0.1097   0.0016   1.0000
<< Back to WORTMANN FX 60-126/1 AIRFOIL (fx601261-il)

Polar data table (+)

Polar graphs


<< Back to WORTMANN FX 60-126/1 AIRFOIL (fx601261-il)