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

GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 562 AIRFOIL (goe562-il)
Reynolds number: 200,000
Max Cl/Cd: 70.55 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe562-il-200000-n5.txt
Download as CSV file: xf-goe562-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 562 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.1784   0.12353   0.12016  -0.0519   0.9880   0.0258
 -11.000  -0.1704   0.11956   0.11620  -0.0553   0.9800   0.0259
 -10.750  -0.1808   0.11758   0.11431  -0.0519   0.9341   0.0260
 -10.500  -0.1625   0.11211   0.10879  -0.0592   0.9130   0.0261
 -10.250  -0.1373   0.10577   0.10236  -0.0695   0.8970   0.0262
 -10.000  -0.0832   0.10009   0.09652  -0.0766   0.8791   0.0277
  -9.750  -0.0385   0.09385   0.09010  -0.0891   0.8578   0.0290
  -9.500  -0.0008   0.08728   0.08333  -0.1048   0.8322   0.0319
  -9.000   0.0473   0.07658   0.07225  -0.1240   0.7745   0.0323
  -8.250   0.0746   0.06801   0.06331  -0.1266   0.7105   0.0304
  -8.000   0.0716   0.06490   0.06013  -0.1278   0.6962   0.0299
  -7.750   0.0643   0.06194   0.05713  -0.1279   0.6843   0.0291
  -7.500   0.0566   0.05887   0.05401  -0.1278   0.6741   0.0288
  -7.250   0.0543   0.05581   0.05084  -0.1279   0.6649   0.0292
  -7.000   0.0538   0.05266   0.04759  -0.1278   0.6566   0.0301
  -6.750   0.0536   0.04927   0.04404  -0.1271   0.6489   0.0304
  -6.500   0.0542   0.04565   0.04024  -0.1259   0.6421   0.0309
  -6.250   0.0567   0.04232   0.03669  -0.1241   0.6353   0.0309
  -6.000   0.0601   0.03898   0.03306  -0.1217   0.6295   0.0311
  -5.750   0.0577   0.03454   0.02809  -0.1176   0.6242   0.0327
  -5.500   0.0733   0.03366   0.02717  -0.1163   0.6178   0.0335
  -5.250   0.0863   0.03212   0.02544  -0.1143   0.6120   0.0341
  -5.000   0.0989   0.03039   0.02348  -0.1119   0.6060   0.0346
  -4.750   0.1120   0.02861   0.02138  -0.1094   0.6007   0.0349
  -4.500   0.1279   0.02735   0.01990  -0.1074   0.5951   0.0362
  -4.250   0.1437   0.02593   0.01818  -0.1051   0.5896   0.0373
  -4.000   0.1605   0.02449   0.01640  -0.1028   0.5848   0.0377
  -3.750   0.1790   0.02324   0.01485  -0.1009   0.5801   0.0381
  -3.500   0.1987   0.02216   0.01350  -0.0991   0.5748   0.0385
  -3.250   0.2200   0.02149   0.01252  -0.0976   0.5699   0.0397
  -3.000   0.2419   0.02093   0.01169  -0.0962   0.5650   0.0404
  -2.750   0.2643   0.01984   0.01045  -0.0951   0.5598   0.0408
  -2.500   0.2879   0.01909   0.00955  -0.0942   0.5551   0.0413
  -2.250   0.3125   0.01851   0.00883  -0.0936   0.5512   0.0419
  -2.000   0.3367   0.01800   0.00827  -0.0928   0.5463   0.0423
  -1.750   0.3609   0.01757   0.00777  -0.0920   0.5415   0.0431
  -1.500   0.3851   0.01729   0.00741  -0.0913   0.5372   0.0448
  -1.250   0.4094   0.01698   0.00705  -0.0905   0.5329   0.0461
  -1.000   0.4331   0.01667   0.00671  -0.0896   0.5281   0.0470
  -0.750   0.4562   0.01639   0.00637  -0.0886   0.5234   0.0479
  -0.500   0.4795   0.01618   0.00608  -0.0877   0.5193   0.0488
  -0.250   0.5016   0.01598   0.00588  -0.0865   0.5144   0.0497
   0.000   0.5228   0.01577   0.00566  -0.0851   0.5090   0.0506
   0.250   0.5442   0.01557   0.00540  -0.0838   0.5047   0.0525
   0.500   0.5657   0.01546   0.00530  -0.0825   0.4994   0.0554
   0.750   0.5876   0.01540   0.00520  -0.0813   0.4947   0.0593
   1.000   0.6101   0.01536   0.00509  -0.0802   0.4907   0.0627
   1.250   0.6319   0.01530   0.00500  -0.0789   0.4861   0.0692
   1.500   0.6527   0.01522   0.00498  -0.0775   0.4806   0.0854
   2.000   0.8994   0.01423   0.00586  -0.1208   0.4608   1.0000
   2.500   0.9423   0.01452   0.00605  -0.1184   0.4529   1.0000
   2.750   0.9634   0.01467   0.00616  -0.1171   0.4486   1.0000
   3.000   0.9840   0.01483   0.00628  -0.1158   0.4445   1.0000
   3.250   1.0048   0.01500   0.00639  -0.1145   0.4409   1.0000
   3.500   1.0256   0.01516   0.00656  -0.1132   0.4363   1.0000
   3.750   1.0459   0.01533   0.00672  -0.1118   0.4318   1.0000
   4.000   1.0657   0.01552   0.00687  -0.1103   0.4279   1.0000
   4.250   1.0857   0.01571   0.00705  -0.1089   0.4236   1.0000
   4.500   1.1054   0.01589   0.00726  -0.1074   0.4189   1.0000
   4.750   1.1242   0.01609   0.00745  -0.1058   0.4143   1.0000
   5.000   1.1427   0.01631   0.00764  -0.1041   0.4100   1.0000
   5.250   1.1613   0.01651   0.00789  -0.1024   0.4048   1.0000
   5.500   1.1785   0.01672   0.00812  -0.1005   0.3992   1.0000
   5.750   1.1952   0.01697   0.00834  -0.0985   0.3946   1.0000
   6.000   1.2121   0.01718   0.00861  -0.0965   0.3889   1.0000
   6.250   1.2262   0.01742   0.00885  -0.0940   0.3833   1.0000
   6.500   1.2401   0.01766   0.00912  -0.0914   0.3784   1.0000
   6.750   1.2542   0.01790   0.00941  -0.0890   0.3722   1.0000
   7.000   1.2666   0.01818   0.00969  -0.0862   0.3666   1.0000
   7.250   1.2806   0.01846   0.01002  -0.0838   0.3606   1.0000
   7.500   1.2932   0.01878   0.01038  -0.0812   0.3542   1.0000
   7.750   1.3054   0.01913   0.01075  -0.0785   0.3483   1.0000
   8.000   1.3183   0.01949   0.01117  -0.0761   0.3417   1.0000
   8.250   1.3285   0.01993   0.01161  -0.0732   0.3352   1.0000
   8.500   1.3399   0.02037   0.01210  -0.0707   0.3272   1.0000
   8.750   1.3481   0.02092   0.01266  -0.0677   0.3187   1.0000
   9.000   1.3551   0.02154   0.01329  -0.0646   0.3092   1.0000
   9.250   1.3626   0.02222   0.01398  -0.0618   0.3000   1.0000
   9.500   1.3657   0.02310   0.01484  -0.0585   0.2892   1.0000
   9.750   1.3713   0.02399   0.01574  -0.0557   0.2814   1.0000
  10.000   1.3745   0.02505   0.01682  -0.0529   0.2720   1.0000
  10.250   1.3773   0.02624   0.01801  -0.0502   0.2647   1.0000
  10.500   1.3821   0.02741   0.01923  -0.0480   0.2581   1.0000
  10.750   1.3849   0.02879   0.02064  -0.0457   0.2510   1.0000
  11.000   1.3864   0.03034   0.02222  -0.0436   0.2444   1.0000
  11.250   1.3905   0.03181   0.02377  -0.0419   0.2381   1.0000
  11.500   1.3904   0.03367   0.02565  -0.0401   0.2324   1.0000
  11.750   1.3922   0.03549   0.02756  -0.0386   0.2248   1.0000
  12.000   1.3884   0.03785   0.02994  -0.0370   0.2170   1.0000
  12.250   1.3893   0.03995   0.03213  -0.0358   0.2088   1.0000
  12.500   1.3831   0.04276   0.03496  -0.0346   0.1987   1.0000
  12.750   1.3787   0.04553   0.03778  -0.0336   0.1843   1.0000
  13.000   1.3559   0.05026   0.04236  -0.0324   0.1378   1.0000
  13.250   1.3323   0.05524   0.04720  -0.0314   0.1138   1.0000
  13.500   1.2979   0.06162   0.05343  -0.0305   0.0783   1.0000
  13.750   1.2801   0.06641   0.05822  -0.0301   0.0637   1.0000
  14.000   1.2641   0.07117   0.06300  -0.0299   0.0524   1.0000
  14.250   1.2493   0.07598   0.06787  -0.0300   0.0401   1.0000
  14.500   1.2353   0.08082   0.07275  -0.0303   0.0287   1.0000
  14.750   1.2220   0.08570   0.07770  -0.0307   0.0243   1.0000
  15.000   1.2113   0.09033   0.08241  -0.0311   0.0227   1.0000
<< Back to GOE 562 AIRFOIL (goe562-il)

Polar data table (+)

Polar graphs


<< Back to GOE 562 AIRFOIL (goe562-il)