NACA 5-H-10 AIRFOIL (n5h10-il)
NACA 5-H-10 AIRFOIL - NACA 5-H-10 rotorcraft airfoil
Details | Dat file | Parser | |
(n5h10-il) NACA 5-H-10 AIRFOIL NACA 5-H-10 rotorcraft airfoil Max thickness 9.9% at 40% chord. Max camber 2.3% at 40% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
NACA 5-H-10 AIRFOIL 18. 18. 0.0000000 0.0000000 0.0125000 0.0150000 0.0250000 0.0210000 0.0500000 0.0310000 0.0750000 0.0380000 0.1000000 0.0437000 0.1500000 0.0530000 0.2000000 0.0610000 0.2500000 0.0660000 0.3000000 0.0690000 0.4000000 0.0720000 0.5000000 0.0670000 0.6000000 0.0550000 0.7000000 0.0390000 0.8000000 0.0200000 0.9000000 0.0067000 0.9600000 0.0027000 1.0000000 0.0000000 0.0000000 0.0000000 0.0125000 -.0093000 0.0250000 -.0120000 0.0500000 -.0150000 0.0750000 -.0168000 0.1000000 -.0185000 0.1500000 -.0210000 0.2000000 -.0230000 0.2500000 -.0240000 0.3000000 -.0250000 0.4000000 -.0270000 0.5000000 -.0280000 0.6000000 -.0290000 0.7000000 -.0290000 0.8000000 -.0240000 0.9000000 -.0110000 0.9600000 -.0047000 1.0000000 0.0000000 |
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Similar airfoils
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Polars for NACA 5-H-10 AIRFOIL (n5h10-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
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n5h10-il | 50,000 | 9 | 30.8 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n5h10-il | 50,000 | 5 | 31.9 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n5h10-il | 100,000 | 9 | 48.4 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n5h10-il | 100,000 | 5 | 48.9 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n5h10-il | 200,000 | 9 | 65.8 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n5h10-il | 200,000 | 5 | 56.2 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n5h10-il | 500,000 | 9 | 70 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n5h10-il | 500,000 | 5 | 58.7 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
n5h10-il | 1,000,000 | 9 | 74.9 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
n5h10-il | 1,000,000 | 5 | 69 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |