• Steel and Composite Structures
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• v.46 no.5
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• pp.611-619
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• 2023

In the present work, the flutter characteristics of porous nanocomposite cylindrical shells, reinforced with graphene platelets (GPLs) in supersonic airflow, have been investigated. Different distributions for GPLs and porosities have been considered which are named uniform and non-uniform distributions thorough the shell's thickness. The effective material properties have been determined via Halpin-Tsai micromechanical model. The cylindrical shell formulation considering supersonic airflow has been developed in the context of first-order shell and first-order piston theories. The governing equations have been solved using Galerkin's method to find the frequency-pressure plots. It will be seen that the flutter points of the shell are dependent on the both amount and distribution of porosities and GPLs and also shell geometrical parameters.

• Proceedings of the Materials Research Society of Korea Conference
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• 2007.05a
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• pp.42-42
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• 2007

• Journal of the Korean Society of Visualization
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• v.9 no.3
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• pp.30-37
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• 2011

The thrust vectoring concept has been used for use in new advanced supersonic aircraft. This study presents the performance characteristics of the thrust vectoring nozzle by visualizing the shock behaviors with Schlieren method. The scaled models were designed and manufactured to see the shock behaviors of the various airflow condition. Also we executed experimental tests to see the geometrical effects of the thrust vector nozzle by changing pitch angle and length of pitch flaps. From this study we could understand the supersonic flow characteristics of the thrust vector nozzle. The total thrust of thrust vector nozzle is diminished by increasing the flap angle. But there is an optimum flap length ratio for attaining the highest thrust level and proper pitch effect.

• Journal of the Korean Society of Combustion
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• v.7 no.3
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• pp.32-36
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• 2002

A numerical study was carried out to investigate combustion phenomena in a model Scramjet engine, which had been experimentally studied at the University of Tokyo using a high-enthalpy supersonic wind tunnel. The main airflow was Mach number 2.0 and the total temperature of hot flow was 1800K. Equivalence ratio was set to be 0.26 which is higher than that of experiment to investigate the effect of strong precombustion shock. The results showed that self-ignition occurred at the rear bottom wall of the combustor and combined with the shear layer flame between fuel jet and main airflow. Then, precombustion shock was generated at the step location and reversely enhanced the mixing and combustion process behind the shock. Due to the high equivalence ratio, the precombustion shock moved upstream of the step compared with that of experiment.

• 한국연소학회:학술대회논문집
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• 2001.06a
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• pp.127-132
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• 2001

A numerical study was carried out to investigate the combustion phenomena in a model Scramjet engine, which had been experimentally studied in the University of Tokyo using a high-enthalpy supersonic wind tunnel. The main airflow was 2.0 in Mach number and the total temperature of hot flow was 1800K. Equivalence ratio was set to be rather higher value of 0.26 than that of experiment to investigate the effect of strong precombustion shock. The results showed that self-ignition occurred at the rear bottom wall of the combustor and combined with the shear layer flame between fuel jet and main airflow. Then, precombustion shock was generated at the step location and reversely enhanced the mixing and combustion process behind the shock. Due to the high equivalence ratio, the precombustion shock moved upstream of the step compared with that of experiment.

• Advances in aircraft and spacecraft science
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• v.5 no.4
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• pp.445-458
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• 2018

Turbulent airflow in channels of rectangular cross section with symmetric centerbodies is studied numerically. Shock wave configurations formed in the channel and in front of the entrance are examined. Solutions of the unsteady Reynolds-averaged Navier-Stokes equations are obtained with finite-volume solvers of second-order accuracy. The solutions demonstrate an expulsion/swallowing of the shocks with variations of the free-stream Mach number or angle of attack. Effects of the centerbody length and thickness on the shock wave stability and flow bifurcation are examined. Bands of the Mach number and angle of attack, in which there exist non-unique flow fields, are identified.

• Structural Engineering and Mechanics
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• v.84 no.4
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• pp.437-452
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• 2022

For the first time, the higher-order shear and normal deformable plate theory (HOSNDPT) is used for the vibration and flutter analyses of the multilayer functionally graded graphene platelets reinforced composite (FG-GPLRC) plates under supersonic airflow. For modeling the supersonic airflow, the linear piston theory is adopted. In HOSNDPT, Legendre polynomials are used to approximate the components of the displacement field in the thickness direction. So, all stress and strain components are encountered. Either uniform or three kinds of non-uniform distribution of graphene platelets (GPLs) into polymer matrix are considered. The Young modulus of the FG-GPLRC plate is estimated by the modified Halpin-Tsai model, while the Poisson ratio and mass density are determined by the rule of mixtures. The Hamilton's principle is used to obtain the governing equations of motion and the associated boundary conditions of the plate. For solving the plate's equations of motion, the Galerkin approach is applied. A comparison for the natural frequencies obtained based on the present investigation and those of three-dimensional elasticity theory shows a very good agreement. The flutter boundaries for FG-GPLRC plates based on HOSNDPT are described and the effects of GPL distribution patterns, the geometrical parameters and the weight fraction of GPLs on the flutter frequencies and flutter aerodynamic pressure of the plate are studied in detail. The obtained results show that by increasing 0.5% of GPLs into polymer matrix, the flutter aerodynamic pressure increases approximately 117%, 145%, 166% and 196% for FG-O, FG-A, UD and FG-X distribution patterns, respectively.

• Composites Research
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• v.13 no.5
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• pp.50-59
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• 2000

In this paper, two methods to suppress flutter of the composite panel are examined. First, in the active control method, a controller based on the linear optimal control theory is designed and control input voltage is applied on the actuators and a PZT is used as actuator. Second, a new technique, passive suppression scheme, is suggested for suppression of the nonlinear panel flutter. In the passive suppression scheme, a shunt circuit which consists of inductor-resistor is used to increase damping of the system and as a result the flutter can be attenuated. A passive damping technology, which is believed to be more robust suppression system in practical operation, requires very little or no electrical power and additional apparatuses such as sensor system and controller are not needed. To achieve the great actuating force/damping effect, the optimal shape and location of the actuators are determined by using genetic algorithms. The governing equations are derived by using extended Hamilton's principle. They are based on the nonlinear von Karman strain-displacement relationship for the panel structure and quasi-steady first-order piston theory for the supersonic airflow. The discretized finite element equations are obtained by using 4-node conforming plate element. A modal reduction is performed to the finite element equations in order to suppress the panel flutter effectively and nonlinear-coupled modal equations are obtained. Numerical suppression results, which are based on the reduced nonlinear modal equations, are presented in time domain by using Newmark nonlinear time integration method.