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

CHEN AIRFOIL (chen-il)

CHEN AIRFOIL - University of Illinois at Urbana-Champaign / Chen high lift airfoil


Airfoil chen-il
Details Dat file Parser  
(chen-il) CHEN AIRFOIL
University of Illinois at Urbana-Champaign / Chen high lift airfoil
Max thickness 12.5% at 27.7% chord.
Max camber 7.8% at 25.4% chord
Source UIUC Airfoil Coordinates Database
Source dat file
The dat file is in Lednicer format
 CHEN AIRFOIL
       51.       48.

 0.0000000 0.0000000
 0.0003558 0.0017755
 0.0018436 0.0084703
 0.0047082 0.0160470
 0.0090325 0.0242980
 0.0147986 0.0330236
 0.0220880 0.0420161
 0.0304777 0.0510117
 0.0407061 0.0604470
 0.0522330 0.0698676
 0.0651491 0.0791655
 0.0793468 0.0882499
 0.0949072 0.0969123
 0.1115329 0.1051795
 0.1293059 0.1128435
 0.1482170 0.1198041
 0.1679599 0.1259887
 0.1885260 0.1312975
 0.2098976 0.1355311
 0.2319581 0.1384987
 0.2543924 0.1400281
 0.2771567 0.1396213
 0.2997316 0.1359180
 0.3245689 0.1295050
 0.3532844 0.1228496
 0.3839673 0.1158192
 0.4161386 0.1086572
 0.4493841 0.1011996
 0.4836304 0.0937538
 0.5185713 0.0862464
 0.5540261 0.0788946
 0.5895069 0.0718414
 0.6249059 0.0649964
 0.6599432 0.0585850
 0.6943033 0.0524342
 0.7278053 0.0467607
 0.7601339 0.0413920
 0.7911083 0.0365444
 0.8205041 0.0319369
 0.8482387 0.0277774
 0.8739075 0.0240015
 0.8976089 0.0206002
 0.9190280 0.0174009
 0.9380744 0.0145117
 0.9546400 0.0118420
 0.9688153 0.0092833
 0.9802941 0.0067623
 0.9890499 0.0039802
 0.9954385 0.0027124
 0.9987892 0.0010106
 1.0000000 0.0000000

 0.0000000 0.0000000
 0.0001465 -.0040284
 0.0011596 -.0084347
 0.0031979 -.0114257
 0.0061967 -.0136988
 0.0106477 -.0153975
 0.0175501 -.0165099
 0.0265157 -.0169012
 0.0375885 -.0160736
 0.0517692 -.0140150
 0.0690841 -.0104265
 0.0864429 -.0063401
 0.1068134 -.0021185
 0.1271930 0.0022028
 0.1496093 0.0069462
 0.1759774 0.0111393
 0.2062535 0.0142842
 0.2263533 0.0154178
 0.2663509 0.0153939
 0.2962248 0.0139564
 0.3260373 0.0118217
 0.3587833 0.0089256
 0.3885173 0.0058943
 0.4177550 0.0029070
 0.4567609 -.0004423
 0.4935460 -.0028925
 0.5295632 -.0049737
 0.5648774 -.0070930
 0.5994238 -.0088433
 0.6332201 -.0100253
 0.6662313 -.0110375
 0.6983493 -.0119707
 0.7297727 -.0128425
 0.7600055 -.0136089
 0.7888490 -.0142526
 0.8164113 -.0146824
 0.8425377 -.0143830
 0.8669041 -.0136268
 0.8899892 -.0126568
 0.9114051 -.0113384
 0.9306377 -.0098270
 0.9479849 -.0081488
 0.9630589 -.0061692
 0.9756256 -.0042691
 0.9856419 -.0029464
 0.9930255 -.0019932
 0.9980277 -.0008293
 1.0000000 0.0000000
No parser warnings Send to airfoil plotter
Add to comparison
Lednicer format dat file
Selig format dat file

Similar airfoils

NACA M27 AIRFOILPreviewDetails
GOE 630 AIRFOILPreviewDetails
GOE 447 AIRFOILPreviewDetails
NACA M24 AIRFOILPreviewDetails
GOE 366 AIRFOILPreviewDetails
GOE 320 (HANSA-BRANDENBURG II.1)PreviewDetails
NACA M21 AIRFOILPreviewDetails
GOE 322 (HANSA-BRANDENBURG IV.1)PreviewDetails
EPPLER 422 AIRFOILPreviewDetails
FX 74-CL6-140PreviewDetails

Polars for CHEN AIRFOIL (chen-il)

PlotAirfoilReynolds #NcritMax Cl/CdDescriptionSource 
   chen-il50,00094.2 at α=9°Mach=0 Ncrit=9Xfoil predictionDetails
   chen-il50,00054.8 at α=11.5°Mach=0 Ncrit=5Xfoil predictionDetails
   chen-il100,00094.9 at α=10.5°Mach=0 Ncrit=9Xfoil predictionDetails
   chen-il100,00058.7 at α=2.75°Mach=0 Ncrit=5Xfoil predictionDetails
   chen-il200,00095.6 at α=11.25°Mach=0 Ncrit=9Xfoil predictionDetails
   chen-il200,000519.6 at α=5.5°Mach=0 Ncrit=5Xfoil predictionDetails
   chen-il500,000913.6 at α=3.25°Mach=0 Ncrit=9Xfoil predictionDetails
   chen-il500,000556.5 at α=12°Mach=0 Ncrit=5Xfoil predictionDetails
   chen-il1,000,0009125.4 at α=14.75°Mach=0 Ncrit=9Xfoil predictionDetails
   chen-il1,000,0005123 at α=13°Mach=0 Ncrit=5Xfoil predictionDetails
Reynolds number calculator
Set Reynolds number and Ncrit rangeLowHigh
Reynolds Number
NCrit