NI\S/\. National Aeronautics and. Space Administration. 1111111111111 IIII 11111 IIIII 11111 11111 IIII 1111. NFOI053 !IBRARY ~~PV. LANGLEY RES~ARCH ...
USAAVSCOM Technical Report 85-A-3
NASA Technical Memorandum 86682
NASA-TM-86682
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Finite - Difference Computations of Rotor Loads F.X. Caradonna and C. Tung
April 1985
!IBRARY ~~PV LANGLEY RES~ARCH CENTER Ll8.~ARY, NASA
Ho.:'.1PTON, VIRGiNIA
NI\S/\
National Aeronautics and Space Administration
1111111111111 IIII 11111 IIIII 11111 11111 IIII 1111
NFOI053
United States Army Aviation Systems Command
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NASA Technical Memorandum 86682
USAAVSCOM Technical Report 85-A-3
-.
Finite - Difference Computations of Rotor Loads F. X. Caradonna C. Tung, Aeromechanics Laboratory, U. S. Army Research and Technology Laboratories (AVSCOM) Ames Research Center, Moffett Field, California
April 1985
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NI\S/\
National Aeronautics and Space Administration
Ames Research Center Moffett Field, California 94035
United States Army Aviation Systems Command St. Louis, Missouri 63120
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FINITE-DIFFERENCE COMPUTATIONS OF ROTOR LOADS F. X. Caradonna and C. Tung Aeromechanics Laboratory, U.S. Army Research and Technology Laboratories--AVSCOM NASA Ames Research Center Moffett Field, California
"
Abstract This paper demonstrates the current and future potential of finite-difference methods for solving real rotor problems which now rely largely on empiricism. The demonstration consists of a simple means of combining existing finitedifference, integral, and comprehensive loads codes to predict real transonic rotor flows. These computations are performed for hover and high-advance-ratio flight. Comparisons are made with experimental pressure data.
x
• x'/c, nondimensional streamwise coordinate (prime denotes a physical quantity)
y
• y'/R, nondimensional spanwise coordinate
(prime denotes a physical quantity)
z
• z'/C6'/3, scaled nondimensional normal
coordinate (prime denotes a physical quantity) • angle of attack, deg • partial angle of attack, deg
Notation A
2 2/3 • M2/AR 6 T
AR
- Ric, aspect ratio
B
• 2M2f/AR6 2/3 T
C
2 2/3 • 1/AR 6
y
= specific heat ratio
6
• thickness ratio collective pitch angle, deg
Sc
S,c • longitudinal cyclic input, deg S1S • lateral cyclic input, deg
CL • lift coefficient
• advance ratio
CT • thrust coefficient
2/3
¢'/ORco , a scaled velocity perturbation potential
chord
c
~
D
• Bg
