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Zhang X.F.,NRC Institute for Aerospace Research | Hodson H.,University of Cambridge
Journal of Turbomachinery | Year: 2010

The effects of Reynolds numbers and the freestream turbulence intensities (FSTIs) on the unsteady boundary layer development on an ultra-high-lift low-pressure turbine airfoil, so-called T106C, are investigated. The measurements were carried out at both Tu =0.5% and 4.0% within a range of Reynolds numbers, based on the blade chord and the isentropic exit velocity, between 100,000 and 260,000. The interaction between the unsteady wake and the boundary layer depends on both the strength of the wake and the status of the boundary layer. At Tu=0.5%, both the wake's high turbulence and the negative jet behavior of the wake dominate the interaction between the unsteady wake and the separated boundary layer on the suction surface of the airfoil. Since the wake turbulence cannot induce transition before separation on this ultra-high-lift blade, the negative jet of the wake has the opportunity to induce a rollup vortex. At Tu=4.0%, the time-mean separation on the suction surface is much smaller. With elevated FSTI, the turbulence in the wake just above the boundary layer is no longer distinguishable from the background turbulence level. The unsteady boundary layer transition is dominated by the wake's negative jet induced boundary layer variation. © 2010 by ASME.

Poirel D.,Royal Military College of Canada | Yuan W.,NRC Institute for Aerospace Research
Journal of Fluids and Structures | Year: 2010

Experimental observations of self-sustained pitch oscillations of a NACA 0012 airfoil at transitional Reynolds numbers were recently reported. The aeroelastic limit cycle oscillations, herein labelled as laminar separation flutter, occur in the range 5.0×104≤Rec≤1.3×105. They are well behaved, have a small amplitude and oscillate about O{middle tilde}=0°. It has been speculated that laminar separation leading to the formation of a laminar separation bubble, occurring at these Reynolds numbers, plays an essential role in these oscillations. This paper focuses on the Rec=7.7×104 case, with the elastic axis located at 18.6% chord. Considering that the experimental rig acts as a dynamic balance, the aerodynamic moment is derived and is empirically modelled as a generalized Duffing-van-der-Pol nonlinearity. As expected, it behaves nonlinearly with pitch displacement and rate. It also indicates a dynamically unstable equilibrium point, i.e. negative aerodynamic damping. In addition, large eddy simulations of the flow around the airfoil undergoing prescribed simple harmonic motion, using the same amplitude and frequency as the aeroelastic oscillations, are performed. The comparison between the experiment and simulations is conclusive. Both approaches show that the work done by the airflow on the airfoil is positive and both have the same magnitude. The large eddy simulation (LES) computations indicate that at O{middle tilde}=0°, the pitching motion induces a lag in the separation point on both surfaces of the airfoil resulting in negative pitching moment when pitching down, and positive moment when pitching up, thus feeding the LCO. © 2010.

Liao M.,NRC Institute for Aerospace Research
Engineering Fracture Mechanics | Year: 2010

This paper presents a study on dislocation theory based short/small crack modeling, and its application for short crack growth life analysis on 2024-T351 aluminum specimens. The dislocation theory was applied to determine the crack tip opening displacement (CTOD) of a microstructurally short crack by taking into account the effects of microstructural features, such as grain size, orientation, and grain boundary. The CTOD was then used as the parameter for calculating the short crack growth rate. In this work, an existing CTOD model was modified for estimating the short crack life of 2024-T351 single edge notch tension (SENT) specimens, which were tested in an Advisory Group for Aerospace Research and Development (AGARD) program [Newman JC, Edward PR. Short-crack growth behavior in an aluminum alloy: an AGARD cooperative test programme. AGARD-R-732, NATO, Advisory Group for Aerospace Research and Development; 1988]. The analytical results matched the test results reasonably well. Crown Copyright © 2009.

Li G.,NRC Institute for Aerospace Research
Journal of Mechanics of Materials and Structures | Year: 2010

In this paper, the adhesive stresses in unbalanced bonded single-strap butt joints are theoretically studied. Mathematical difficulties in the analysis of high order differential equations were solved and closedform solutions for both the adhesive peel and shear stresses have been successfully developed. In the proposed solutions the adherends and doublers can be different in material and thickness. Peak stresses are located at the bonded overlap edges, especially at the inner edges. In addition, two-dimensional geometrically nonlinear finite element analyses were carried out to study the adhesive stresses in two different bonded butt joints. One was a special butt joint case with the adherends and doubler of identical material and thickness, and the other was a general butt joint case with different adherends and doubler. Good agreement in the adhesive stresses between the closed-form solutions and finite element results has been achieved. The single-strap butt joint actually consists of two single-lap joints; thus, the adhesive stress solutions can be further applied to unbalanced single-lap joints.

Ghinet S.,NRC Institute for Aerospace Research | Atalla N.,Universite de Sherbrooke
Computers and Structures | Year: 2011

The paper describes the modeling of thick composite laminate and sandwich plates and beams with linear viscoelastic treatments. A discrete laminate model (DLM) is described, validated and compared to numerical spectral finite elements method (SFEM), finite element method (FEM) and experimental results. The DLM approach assumes each layer as thick laminate with orthotropic orientation, rotational inertia and transversal shearing, membrane and bending deformations. First order shear deformation theory is used. The equation of motion is developed following a wave approach based on discrete layer description. It handles symmetrical and asymmetrical layouts of unlimited number of transversal incompressible layers. Next, dilatational (symmetric mode) motion along the core's thickness is considered to complete the DLM solution when applied to the case of symmetric sandwich structures with soft and thick core. The model is compared to a second approach employing spectral finite elements. The latter handles composite laminated plates and beams with orthotropic orientation. It is shown that both approaches estimate accurately the propagating wave solutions of laminated structures. Using these solutions, the input mobility and the mechanical impedance are computed. The two models are successfully compared to classical finite elements as well as to experimental results for different boundary conditions. Moreover, the equivalent damping loss factor of composite laminate plates with viscoelastic treatment is addressed and the influence of the heading direction is discussed. © 2011 Elsevier Ltd. All rights reserved.

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