Finite-difference computations of rotor loads

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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

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NI\S/\

National Aeronautics and Space Administration

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United States Army Aviation Systems Command

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NASA Technical Memorandum 86682

USAAVSCOM Technical Report 85-A-3

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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

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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