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GOE 280 (DAIMLER XI) AIRFOIL (goe280-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 280 (DAIMLER XI) AIRFOIL (goe280-il)
Reynolds number: 500,000
Max Cl/Cd: 91.58 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe280-il-500000.txt
Download as CSV file: xf-goe280-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 280 (DAIMLER XI) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.750   0.5479   0.01052   0.00346  -0.0822   0.5960   0.0461
   1.000   0.5754   0.01038   0.00318  -0.0821   0.5657   0.0460
   1.250   0.6032   0.01031   0.00299  -0.0821   0.5416   0.0462
   1.500   0.6309   0.01031   0.00288  -0.0821   0.5212   0.0473
   1.750   0.6589   0.01031   0.00280  -0.0822   0.5027   0.0480
   2.000   0.6869   0.01031   0.00272  -0.0823   0.4858   0.0480
   2.250   0.7148   0.01035   0.00269  -0.0823   0.4687   0.0485
   2.500   0.7425   0.01044   0.00268  -0.0824   0.4487   0.0496
   2.750   0.7702   0.01054   0.00267  -0.0824   0.4266   0.0564
   3.000   0.7907   0.00876   0.00284  -0.0813   0.4074   1.0000
   3.250   0.8182   0.00897   0.00293  -0.0813   0.3859   1.0000
   3.500   0.8453   0.00923   0.00304  -0.0813   0.3636   1.0000
   3.750   0.8721   0.00953   0.00319  -0.0812   0.3357   1.0000
   4.000   0.8979   0.00997   0.00339  -0.0811   0.2942   1.0000
   4.250   0.9233   0.01048   0.00366  -0.0810   0.2508   1.0000
   4.500   0.9490   0.01094   0.00392  -0.0808   0.2179   1.0000
   4.750   0.9741   0.01148   0.00423  -0.0806   0.1825   1.0000
   5.000   0.9994   0.01196   0.00453  -0.0805   0.1548   1.0000
   5.250   1.0250   0.01238   0.00484  -0.0803   0.1355   1.0000
   5.500   1.0504   0.01281   0.00514  -0.0802   0.1187   1.0000
   5.750   1.0747   0.01338   0.00550  -0.0799   0.0901   1.0000
   6.000   1.0958   0.01443   0.00621  -0.0791   0.0389   1.0000
   6.250   1.1194   0.01509   0.00672  -0.0786   0.0181   1.0000
   6.500   1.1449   0.01543   0.00715  -0.0783   0.0163   1.0000
   6.750   1.1696   0.01589   0.00772  -0.0778   0.0151   1.0000
   7.000   1.1937   0.01641   0.00835  -0.0773   0.0143   1.0000
   7.250   1.2171   0.01703   0.00908  -0.0767   0.0137   1.0000
   7.500   1.2395   0.01775   0.00994  -0.0759   0.0132   1.0000
   7.750   1.2608   0.01858   0.01089  -0.0749   0.0128   1.0000
   8.000   1.2786   0.01978   0.01223  -0.0735   0.0124   1.0000
   8.250   1.2972   0.02077   0.01331  -0.0723   0.0118   1.0000
   8.500   1.3154   0.02172   0.01435  -0.0709   0.0116   1.0000
   8.750   1.3315   0.02282   0.01555  -0.0693   0.0115   1.0000
   9.000   1.3454   0.02403   0.01685  -0.0673   0.0114   1.0000
   9.250   1.3564   0.02541   0.01831  -0.0649   0.0115   1.0000
   9.500   1.3619   0.02714   0.02011  -0.0618   0.0117   1.0000
   9.750   1.2715   0.01642   0.00983  -0.0455   0.0123   1.0000
  10.750   1.3499   0.01702   0.01103  -0.0407   0.0199   1.0000
  11.000   1.3436   0.01927   0.01341  -0.0376   0.0196   1.0000
  11.250   1.3339   0.02219   0.01643  -0.0347   0.0190   1.0000
  11.500   1.3264   0.02533   0.01967  -0.0323   0.0186   1.0000
  11.750   1.3184   0.02892   0.02334  -0.0301   0.0182   1.0000
  12.000   1.3133   0.03286   0.02731  -0.0280   0.0176   1.0000
  12.250   1.3000   0.03802   0.03266  -0.0259   0.0162   1.0000
  12.500   1.2940   0.04162   0.03641  -0.0250   0.0158   1.0000
  12.750   1.2878   0.04556   0.04049  -0.0241   0.0154   1.0000
  13.000   1.2810   0.04972   0.04478  -0.0233   0.0150   1.0000
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