• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2000.06a
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• pp.1198-1203
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• 2000

In this study, a nonlinear aeroelastic analysis system for the fighter wing with tip-store has been developed additionally in the transonic and supersonic flow region. The unsteady CFD code based on the transonic small disturbance theory has been incorporated to consider the numerical capability for the aerodynamic nonlinear effects. The coupled time-integration method is used to observe the detailed nonlinear aeroelastic responses for elastic wings in their flight. condition. A conservative wing-box model of a fighter wing with tip-store is modeled by MSC/PATRAN and the corresponding free vibration analysis has been performed by MSC/NASTRAN. The results of flutter analyses are presented in the subsonic, transonic and supersonic flow regime.

• Advances in aircraft and spacecraft science
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• v.4 no.4
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• pp.353-369
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• 2017

Flutter is a dangerous phenomenon encountered in flexible structures subjected to aerodynamic forces. This includes aircraft, buildings and bridges. Flutter occurs as a result of interactions between aerodynamic, stiffness, and inertia forces on a structure. In an aircraft, as the speed of the flow increases, there may be a point at which the structural damping is insufficient to damp out the motion which is increasing due to aerodynamic energy being added to the structure. This vibration can cause structural failure, and therefore considering flutter characteristics is an essential part of designing an aircraft. Scientists and engineers studied flutter and developed theories and mathematical tools to analyze the phenomenon. Strip theory aerodynamics, beam structural models, unsteady lifting surface methods (e.g., Doublet-Lattice) and finite element models expanded analysis capabilities. Periodic Structures have been in the focus of research for their useful characteristics and ability to attenuate vibration in frequency bands called "stop-bands". A periodic structure consists of cells which differ in material or geometry. As vibration waves travel along the structure and face the cell boundaries, some waves pass and some are reflected back, which may cause destructive interference with the succeeding waves. This may reduce the vibration level of the structure, and hence improve its dynamic performance. In this paper, for the first time, we analyze the flutter characteristics of a wing with a periodic change in its sandwich construction. The new technique preserves the external geometry of the wing structure and depends on changing the material of the sandwich core. The periodic analysis and the vibration response characteristics of the model are investigated using a finite element model for the wing. Previous studies investigating the dynamic bending response of a periodic sandwich beam in the absence of flow have shown promising results.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.4
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• pp.295-301
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• 2007

The aeroelastic characteristics of a wing with control surface freeplay are investigated. The transonic small disturbance equation is used for unsteady aerodynamic forces in subsonic/transonic region. The fictitious mass method is introduced to apply a modal approach to nonlinear structural models. Nonlinear aeroelastic time responses are calculated by the coupled time integration method. Using these methods, an efficient aeroelastic analysis is achieved for aerodynamic and structural nonlinearities simultaneously. The effects of the aerodynamic nonlinearity, initial flap amplitude, and freeplay magnitude in aeroelastic characteristics are investigated in this study.

• Journal of the Korea Institute of Military Science and Technology
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• v.23 no.4
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• pp.363-370
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• 2020

In this study, pitching stability derivatives of the conical shell configuration is predicted using commercial CFD code. Unsteady flow analysis with forced harmonic motion of the model is performed using overset mesh. The test is conducted about Basic finner missile configuration. The static and dynamic stability derivatives are good agreement with available experimental data. As the same way, a conical shell is analyzed in Mach number 1.6 and various reduced frequency. The static and dynamic derivatives are obtained from the time-pitching moment coefficient histories in each of four cases of mean angle of attack. The variation of reduced frequency is not affected static and dynamic derivatives. Increasing the mean angle of attack, static derivatives are increased slowly. Comparison of the Cm curves at the steady and unsteady state results shows that the Cm curve including the damping effect is lower than otherwise case, approximately 9-18 %.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.7
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• pp.13-18
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• 2004

This paper describes an efficient and robust optimization method for helicopter rotor airfoil design in forward flight. Navier-Stokes analysis was employed to compute the dynamic response of an airfoil, which simulates the unsteady rotor flow-field in forward flight. The optimization system consists of two categories; Response Surface Method to construct the response surface model based on D-optimal 3-level factorial design, and Genetic Algorithm to obtain the optimum solution of a defined objective function including penalty terms of constraints. The influence of design variables and their interactions on the aerodynamic performance was examined through the optimization process.

• Advances in aircraft and spacecraft science
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• v.10 no.5
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• pp.419-437
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• 2023

This paper presents the study of the effects of rigid-body motion simultaneously with the presence of the effects of temporal variation due to the existence of morphing speed on the aeroelastic stability of the two-stage telescopic wings, and hence this is the main novelty of this study. To this aim, Euler-Bernoulli beam theory is used to model the bending-torsional dynamics of the wing. The aerodynamic loads on the wing in an incompressible flow regime are determined by using Peters' unsteady aerodynamic model. The governing aeroelastic equations are discretized employing a finite element method based on the beam-rod model. The effects of rigid-body motion on the length-based stability of the wing are determined by checking the eigenvalues of system. The obtained results are compared with those available in the literature, and a good agreement is observed. Furthermore, the effects of different parameters of rigid-body such as the mass, radius of gyration, fuselage center of gravity distance from wing elastic axis on the aeroelastic stability are discussed. It is found that some parameters can cause unpredictable changes in the critical length and frequency. Also, paying attention to the fuselage parameters and how they affect stability is very important and will play a significant role in the design.

• Advances in aircraft and spacecraft science
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• v.3 no.4
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• pp.447-470
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• 2016

Limit cycle oscillations (LCO) as well as nonlinear aeroelastic analysis of a swept aircraft wing with cubic restoring moments in the pitch degree of freedom is investigated. The unsteady aerodynamic loading applied on the wing is modeled by using the strip theory. The harmonic balance method is used to calculate the LCO frequency and amplitude for the swept wing. Finally the super and subcritical Hopf bifurcation diagrams are plotted. It is concluded that the type of bifurcation and turning point location is sensitive to the system parameters such as wing geometry and sweep angle.

• Advances in aircraft and spacecraft science
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• v.4 no.2
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• pp.203-218
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• 2017

This paper focuses on the computational study of nonlinear effects of unsteady aerodynamics for non-classical aileron buzz. It aims at a comprehensive investigation of the aileron buzz phenomenon under varying flow parameters using the describing function technique with multiple inputs. The limit cycle oscillatory behavior of an asymmetrical airfoil is studied initially using a CFD-based numerical model and direct time marching. Sharp increases in limit cycle amplitude for varying Mach numbers and angles of attack are investigated. An aerodynamic describing function is developed in order to estimate the variation of limit cycle amplitude and frequency with Mach number and angle of attack directly, without time marching. The describing function results are compared to the amplitudes and frequencies predicted by the CFD calculations for validation purposes. Furthermore, a limited sensitivity analysis is presented to demonstrate the potential of the approach for aeroelastic design.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.3
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• pp.215-223
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• 2009

In this paper, the helicopter aerodynamics is simulated in hovering and forward flight. Also, an overlapped grid technique is applied in this simulation to consider the blade motion and moving effects. The Caradonna & Tung's rotor blade was selected to analyze the unsteady aerodynamics in hovering and non-lift forward flight. Also, the AH-1G rotor blade was selected in forward flight. In forward flight case, the numerical trim was applied to determine the cyclic pitching angles using Newton-Raphson method, and the numerical results were in good agreement with experimental data, especially, the BVI effects were well simulated in advancing side in comparison other numerical results. The governing equation is a three dimensional unsteady Euler equation, and the Riemann invariants condition is used for inflow and outflow at the boundary.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.149-158
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• 2006

A review has been made on what kind of method can be applied to predict the dynamic derivatives of the separated PLF(Payload Fairing) halves of a launch vehicle in consideration of technology and budget. An optimal approach is selected considering the geometric characteristics of the PLF halves, the aerodynamic conditions and the required accuracy. The time history of aerodynamic force/moment coefficients are obtained for the forced harmonic motions by solving the unsteady Euler equations derived with respect to the inertial reference frame. and the dynamic derivatives are deduced by integration of the aerodynamic coefficients for one period. In this research, the dynamic derivatives are presented for 0.6$\leq$ M $\leq$2.0, $-180^{\circ}$ $\leq$$\alpha$ $\leq$$180^{\circ}$ and $-90 ^{\circ}$$\leq$$\beta$$\leq$$90 ^{\circ}$.