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

Prandtl-D tip - NASA Preliminary Research Aerodynamic Design To Lower Drag (prandtl-d-tip-ns)

Prandtl-D tip - NASA Preliminary Research Aerodynamic Design To Lower Drag - NASA Preliminary Research Aerodynamic Design To Lower Drag (PRANDTL-D) glider tip airfoil


Airfoil prandtl-d-tip-ns
Details Dat file Parser  
(prandtl-d-tip-ns) Prandtl-D tip - NASA Preliminary Research Aerodynamic Design To Lower Drag
NASA Preliminary Research Aerodynamic Design To Lower Drag (PRANDTL-D) glider tip airfoil
Max thickness 9.6% at 26.8% chord.
Max camber 0% at 0% chord
Source NASA
Source dat file
The dat file is in Selig format
Prandtl-D tip - NASA Preliminary Research Aerodynamic Design To Lower Drag
1.00000 0.00070
0.96091 0.00428
0.94833 0.00540
0.93571 0.00654
0.92307 0.00769
0.89778 0.00999
0.88515 0.01114
0.84728 0.01455
0.82206 0.01679
0.80944 0.01789
0.79683 0.01898
0.78422 0.02006
0.77160 0.02113
0.73374 0.02428
0.72112 0.02531
0.69587 0.02734
0.68325 0.02834
0.67063 0.02933
0.65801 0.03031
0.62017 0.03315
0.60756 0.03406
0.59495 0.03496
0.56975 0.03670
0.55715 0.03754
0.54455 0.03836
0.50678 0.04067
0.49420 0.04139
0.48162 0.04208
0.46904 0.04274
0.43132 0.04453
0.40620 0.04556
0.38108 0.04644
0.36853 0.04682
0.35599 0.04716
0.34346 0.04745
0.33093 0.04770
0.29342 0.04814
0.26848 0.04816
0.25604 0.04807
0.24362 0.04791
0.23122 0.04767
0.21885 0.04736
0.20652 0.04696
0.15762 0.04434
0.14554 0.04338
0.13355 0.04229
0.12166 0.04103
0.09825 0.03801
0.07556 0.03415
0.06460 0.03186
0.03452 0.02347
0.02619 0.02030
0.01925 0.01718
0.01384 0.01430
0.00681 0.00956
0.00460 0.00763
0.00174 0.00435
0.00028 0.00161
0.00002 0.00038
0.00000 0.00000
0.00002 -0.00038
0.00028 -0.00161
0.00174 -0.00435
0.00460 -0.00763
0.00681 -0.00956
0.01384 -0.01430
0.01925 -0.01718
0.02619 -0.02030
0.03452 -0.02347
0.06460 -0.03186
0.07556 -0.03415
0.09825 -0.03801
0.12166 -0.04103
0.13355 -0.04229
0.14554 -0.04338
0.15762 -0.04434
0.20652 -0.04696
0.21885 -0.04736
0.23122 -0.04767
0.24362 -0.04791
0.25604 -0.04807
0.26848 -0.04816
0.29342 -0.04814
0.33093 -0.04770
0.34346 -0.04745
0.35599 -0.04716
0.36853 -0.04682
0.38108 -0.04644
0.40620 -0.04556
0.43132 -0.04453
0.46904 -0.04274
0.48162 -0.04208
0.49420 -0.04139
0.50678 -0.04067
0.54455 -0.03836
0.55715 -0.03754
0.56975 -0.03670
0.59495 -0.03496
0.60756 -0.03406
0.62017 -0.03315
0.65801 -0.03031
0.67063 -0.02933
0.68325 -0.02834
0.69587 -0.02734
0.72112 -0.02531
0.73374 -0.02428
0.77160 -0.02113
0.78422 -0.02006
0.79683 -0.01898
0.80944 -0.01789
0.82206 -0.01679
0.84728 -0.01455
0.88515 -0.01114
0.89778 -0.00999
0.92307 -0.00769
0.93571 -0.00654
0.94833 -0.00540
0.96091 -0.00428
1.00000 0.00070
No parser warnings Send to airfoil plotter
Add to comparison
Lednicer format dat file
Selig format dat file

Similar airfoils

S9026 (9.5%)PreviewDetails
RAE 100 AIRFOILPreviewDetails
LWK 79-100PreviewDetails
RAF 27 AIRFOILPreviewDetails
WORTMANN FX 76-100PreviewDetails
E230 (9.96%)PreviewDetails
NACA 0010PreviewDetails
RAE 101 AIRFOILPreviewDetails
Joukovsky f=0% t=9%PreviewDetails
S9032 (9%)PreviewDetails

Polars for Prandtl-D tip - NASA Preliminary Research Aerodynamic Design To Lower Drag (prandtl-d-tip-ns)

PlotAirfoilReynolds #NcritMax Cl/CdDescriptionSource 
   prandtl-d-tip-ns50,000925.4 at α=4.75°Mach=0 Ncrit=9Xfoil predictionDetails
   prandtl-d-tip-ns100,000935.8 at α=4.5°Mach=0 Ncrit=9Xfoil predictionDetails
   prandtl-d-tip-ns200,000947.1 at α=5°Mach=0 Ncrit=9Xfoil predictionDetails
   prandtl-d-tip-ns500,000962.7 at α=5.5°Mach=0 Ncrit=9Xfoil predictionDetails
   prandtl-d-tip-ns1,000,000974.6 at α=5.5°Mach=0 Ncrit=9Xfoil predictionDetails
   prandtl-d-tip-ns2,000,000987.3 at α=6.75°Mach=0 Ncrit=9Xfoil predictionDetails
   prandtl-d-tip-ns5,000,0009102.8 at α=8.5°Mach=0 Ncrit=9Xfoil predictionDetails
Reynolds number calculator
Set Reynolds number and Ncrit rangeLowHigh
Reynolds Number
NCrit