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

GOE 244 (MVA PR.4) AIRFOIL (goe244-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 244 (MVA PR.4) AIRFOIL (goe244-il)
Reynolds number: 100,000
Max Cl/Cd: 50.11 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe244-il-100000.txt
Download as CSV file: xf-goe244-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 244 (MVA PR.4) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.000   0.1290   0.11042   0.10594  -0.0882   0.7978   0.1150
  -5.750   0.0950   0.11181   0.10739  -0.0872   0.7886   0.1171
  -5.500   0.0802   0.11071   0.10635  -0.0863   0.7781   0.1180
  -5.250   0.1272   0.10518   0.10078  -0.0853   0.7767   0.1210
  -5.000   0.1355   0.10337   0.09898  -0.0844   0.7698   0.1247
  -4.750   0.1155   0.10353   0.09919  -0.0942   0.7572   0.1312
  -4.500   0.1086   0.10180   0.09751  -0.0865   0.7474   0.1319
  -4.250   0.1307   0.09868   0.09438  -0.0824   0.7423   0.1348
  -3.750   0.1954   0.09058   0.08621  -0.0922   0.7365   0.1491
  -3.500   0.1984   0.09059   0.08623  -0.1017   0.7204   0.1602
  -3.250   0.2217   0.08607   0.08172  -0.0980   0.7187   0.1627
  -3.000   0.2585   0.08245   0.07806  -0.0982   0.7174   0.1698
  -2.750   0.5751   0.04770   0.04142  -0.2033   0.7196   0.1051
  -2.500   0.6607   0.04215   0.03547  -0.2145   0.7194   0.0997
  -2.250   0.7547   0.03765   0.03005  -0.2260   0.7193   0.0958
  -2.000   0.7730   0.03737   0.02962  -0.2249   0.7068   0.0959
  -1.750   0.8356   0.03475   0.02682  -0.2295   0.7045   0.0977
  -1.500   0.8994   0.03233   0.02424  -0.2341   0.7022   0.1036
  -1.250   0.9618   0.02987   0.02180  -0.2385   0.6996   0.1115
  -1.000   0.9810   0.02970   0.02162  -0.2367   0.6862   0.1171
  -0.750   1.0480   0.02737   0.01942  -0.2425   0.6816   0.1492
  -0.500   1.0773   0.02719   0.01975  -0.2426   0.6682   0.3276
  -0.250   1.1225   0.02662   0.01903  -0.2436   0.6615   0.3831
   0.000   1.1369   0.02710   0.01954  -0.2405   0.6467   0.4048
   0.250   1.1587   0.02737   0.01977  -0.2385   0.6336   0.4261
   0.500   1.1944   0.02715   0.01941  -0.2382   0.6240   0.4502
   0.750   1.2149   0.02743   0.01958  -0.2369   0.6090   0.4657
   1.000   1.2366   0.02761   0.01970  -0.2353   0.5961   0.4758
   1.250   1.2759   0.02729   0.01910  -0.2368   0.5860   0.4917
   1.500   1.2948   0.02772   0.01944  -0.2355   0.5726   0.5056
   1.750   1.3207   0.02790   0.01952  -0.2350   0.5626   0.5180
   2.000   1.3475   0.02809   0.01958  -0.2348   0.5526   0.5321
   2.250   1.3747   0.02838   0.01975  -0.2349   0.5439   0.5479
   2.500   1.4000   0.02871   0.01999  -0.2346   0.5353   0.5651
   2.750   1.4330   0.02888   0.01998  -0.2355   0.5287   0.5847
   3.000   1.4491   0.02956   0.02071  -0.2338   0.5207   0.6025
   3.250   1.4793   0.02977   0.02079  -0.2341   0.5145   0.6234
   3.500   1.5036   0.03027   0.02124  -0.2337   0.5085   0.6433
   3.750   1.5211   0.03094   0.02194  -0.2323   0.5020   0.6615
   4.000   1.5508   0.03125   0.02213  -0.2327   0.4967   0.6818
   4.250   1.5841   0.03161   0.02234  -0.2337   0.4924   0.7030
   4.500   1.5938   0.03262   0.02350  -0.2312   0.4873   0.7188
   4.750   1.6132   0.03334   0.02424  -0.2302   0.4828   0.7371
   5.000   1.6417   0.03379   0.02461  -0.2305   0.4788   0.7578
   5.250   1.6795   0.03412   0.02478  -0.2324   0.4754   0.7819
   5.500   1.6904   0.03521   0.02602  -0.2302   0.4718   0.8013
   5.750   1.6971   0.03638   0.02737  -0.2275   0.4679   0.8231
   6.000   1.7101   0.03719   0.02834  -0.2256   0.4642   0.8601
   6.250   1.7360   0.03779   0.02893  -0.2258   0.4607   1.0000
   6.500   1.7781   0.03836   0.02931  -0.2287   0.4576   1.0000
   6.750   1.7995   0.03972   0.03066  -0.2285   0.4548   1.0000
   7.000   1.7898   0.04192   0.03309  -0.2238   0.4517   1.0000
   7.250   1.7785   0.04423   0.03557  -0.2190   0.4487   1.0000
   7.500   1.7726   0.04654   0.03800  -0.2152   0.4457   1.0000
   7.750   1.7857   0.04803   0.03950  -0.2139   0.4427   1.0000
   8.000   1.8424   0.04750   0.03873  -0.2180   0.4393   1.0000
   8.250   1.8452   0.04943   0.04072  -0.2152   0.4356   1.0000
   8.500   1.1220   0.14479   0.13842  -0.1933   0.4182   0.8152
<< Back to GOE 244 (MVA PR.4) AIRFOIL (goe244-il)

Polar data table (+)

Polar graphs


<< Back to GOE 244 (MVA PR.4) AIRFOIL (goe244-il)