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Research Papers

Aeroelastic Stability and Response of Composite Swept Wings in Subsonic Flow Using Indicial Aerodynamics

[+] Author and Article Information
S. A. Sina

e-mail: sasina@ae.sharif.edu

H. Haddadpour

Aerospace Engineering Department,
Sharif University of Technology,
Azadi Avenue, P.O. Box 11365-8639,
Tehran, Iran

1Corresponding author.

Contributed by the Design Engineering Division of ASME for publication in the Journal of Vibration and Acoustics. Manuscript received March 2, 2012; final manuscript received February 3, 2013; published online June 18, 2013. Assoc. Editor: Massimo Ruzzene.

J. Vib. Acoust 135(5), 051019 (Jun 18, 2013) (14 pages) Paper No: VIB-12-1053; doi: 10.1115/1.4023992 History: Received March 02, 2012; Revised February 03, 2013

In this study, the aeroelastic stability and response of an aircraft swept composite wing in subsonic compressible flow are investigated. The composite wing was modeled as an anisotropic thin-walled composite beam with the circumferentially asymmetric stiffness structural configuration to establish proper coupling between bending and torsion. Also, the structural model consists of a number of nonclassical effects, such as transverse shear, material anisotropy, warping inhibition, nonuniform torsional model, and rotary inertia. The finite state form of the unsteady aerodynamic loads have been modeled based on the indicial aerodynamic theory and strip theory in the subsonic compressible flow. Novel Mach dependent exponential approximations of the indicial aerodynamic functions have been developed. The extended Galerkin’s method was used to construct the mass, stiffness, and damping matrices of the nonconservative aeroelastic system. Eigen analysis of the system was performed to obtain the aeroelastic instability (divergence and flutter) boundaries. Also, solving the equations of motion in the time domain leads to the aeroelastic response of wing in different flight speeds. The obtained results are compared with the available results in the literature, which reveals an excellent agreement. The numerical results obtained in this article seek to clarify the effects of geometrical and material couplings and flight Mach number on the aeroelastic instability and response of composite wings in subsonic compressible flow.

Copyright © 2013 by ASME
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References

Figures

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Fig. 1

Schematic description of the wing structure and its cross section

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Fig. 2

Cross section coordinate to define complex cross sections of CAS configuration

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Fig. 3

First four coupled bending-torsion frequencies (rad/s) versus fiber angle (deg) for TWB with CAS configuration

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Fig. 4

Variation of bending a33 (Nm2) and torsion a77 (Nm2) stiffness quantities versus ply angle (deg) for composite wing with CAS configuration

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Fig. 5

Mach number at onset of aeroelastic instability versus ply angle (deg) for composite wing with CAS configuration and a = −0.4, Λ = 0 deg

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Fig. 6

Frequency (rad/s) of aeroelastic instability versus ply angle (deg) for composite wing with CAS configuration and a = −0.4, Λ = 0 deg

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Fig. 7

Variation of bending-torsion a¯37 = a37/(a37)θ = 45 deg and transverse shear-torsion a¯56 = a56/(a56)θ = 45 deg stiffness quantities versus ply angle (deg) for composite wing with CAS configuration

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Fig. 8

Normalized responses in bending for composite wing with CAS configuration and θ = −30 deg, a = −0.5, and Λ = 0 deg in M = 0.5 and M = 0.6

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Fig. 9

Normalized responses in bending for composite wing with CAS configuration and θ = −60 deg, a = −0.5, and Λ = 0 deg in M = 0.5 and M = 0.6

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Fig. 10

Subcritical (M = 0.56), flutter (M = 0.57), and supercritical (M = 0.58) normalized responses in bending for composite wing with CAS configuration and θ = −20 deg, a = −0.35, and Λ = 0 deg

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Fig. 11

Subcritical (M = 0.56), flutter (M = 0.57), and supercritical (M = 0.58) normalized responses in torsion for composite wing with CAS configuration and θ = −20 deg, a = −0.35, and Λ = 0 deg

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Fig. 12

Normalized responses in bending for composite wing with CAS configuration and a = −0.45, M = 0.62, and Λ = 0 deg versus a different ply angle

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Fig. 13

Modal frequencies and dampings of the aeroelastic modes plotted versus airspeed for composite TWB with CAS configuration and a = −0.45, Λ = 0 deg, and θ = −70 deg

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Fig. 14

Mach number at onset of the aeroelastic instability versus sweep angle (deg) for composite wing with CAS configuration and a = −0.4, θ = −20 deg, and −30 deg

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Fig. 15

Frequency (rad/s) of aeroelastic instability versus sweep angle (deg) for composite wing with CAS configuration and a = −0.4, θ = −20 deg, and −30 deg

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Fig. 16

Subcritical (M = 0.55) normalized response in torsion for composite wing with CAS configuration and θ = −20 deg, a = −0.5 in different sweep angles

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Fig. 17

Subcritical (M = 0.49), divergence (M = 0.5), and supercritical (M = 0.51) normalized responses in bending for composite wing with CAS configuration and θ = −20 deg, a = −0.4, and Λ = −20 deg

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Fig. 18

Subcritical (M = 0.49), divergence (M = 0.5), and supercritical (M = 0.51) normalized responses in torsion for composite wing with CAS configuration and θ = −20 deg, a = −0.4, and Λ = −20 deg

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Fig. 19

Normalized responses in bending for composite wing with CAS configuration and θ = −25 deg, M = 0.6, and Λ = 0 deg in different a parameter

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