Search

We develop a combined Lagrangian–Eulerian method for transient fluid–structure interaction problems. Based on the ghost fluid framework for improving interface tracking accuracy between a fluid (ho...

Purpose The purpose of this paper is to present the design, analysis and experiments of the active trailing-edge flap (Seoul National University Flap, SNUF) for vibration reduction in the helicopter rotor prior to the small-scale blades planned to test in a whirl tower. Design/methodology/approach The predictions of the hinge moment in both steady and unsteady flows were obtained through computational fluid dynamics calculations. When compared with the results originated from analytical formulations, the proposed method showed improved prediction capabilities. To validate the deflection of the flap under the centrifugal load by rotating, static analysis was conducted using both contact and rotating condition of MSC NASTRAN. The corresponding experiment also was performed using the vertical frame for simulating the effect of the centrifugal force. Findings The hinge moment of the flap is predicted through unsteady analysis in the actuation frequency of 3/rev. The material of the guide in the flap mechanism was selected through static analysis under both contact and rotating condition. Finally, reduction of the deflection occurred because of the load in the axial direction of the hinge like the centrifugal load. Practical implications The important aspects, such as design, analysis, and experiments for the active trailing-edge flap were shown. Originality/value This paper showed the relationship of the displacement, block force and voltage of the piezo-actuator, combined with the hinge moment predicted. The methodology and the experiment were presented for simulating the centrifugal force acting on the flap.

This study presents a three-dimensional nonlinear dynamic formulation based on a co-rotational (CR) approach for solid elements. The CR formulation is relatively efficient, but it is based on the assumption of small strains during large displacement. The novel idea of the present formulation involves the use of the CR formulation through a three-dimensional solid element for inertial quantities in addition to an internal force vector and a stiffness matrix. The present dynamic formulation is derived from Lagrange’s equation of motion. In this procedure, the CR formulation, (i.e., element-independent CR) is one of the most attractive features that is strongly manifested in an efficient manner. Consequently, this obtains the governing equation of motion including motion-driven inertial components (physical quantities induced by the prescribed motion). Four examples are presented to demonstrate the accuracy of the present dynamic formulation. Finally, the results are compared with those obtained by ABAQUS, and the findings reveal that the proposed dynamic formulation is in good agreement with existing predictions.

An improved data-driven non-intrusive model order reduction (MOR) methodology capable of interpolating time-transient flow-fields and other types of data with respect to the parameters is proposed. The proposed MOR method comprises the following two stages: MOR and interpolation. For the MOR, modified proper orthogonal decomposition (POD) is used to collect the parametrically independent POD modes and dependent coefficients. An interpolation of the POD coefficients is conducted through unsupervised machine learning, referred to as the Wasserstein generative adversarial network-gradient penalty (WGAN-GP). By using a deep convolutional neural network, WGAN-GP stabilizes the interpolation across the parameters and ensures an accurate interpolation with few results within the parametric space. An interpolated object is then generated using the parametrically interpolated POD coefficients and relevant independent modes. Next, flow-fields around a stationary cylinder and a plunging airfoil are applied to demonstrate the efficiency and accuracy of the proposed approach, and the influences of the POD modes and parameters on the accuracy are evaluated. Finally, the accuracy and efficiency are compared with those of other methods through the adoption of an accuracy index. Based on the results, the proposed method was found to be effective and efficient for object interpolation.

이 논문은 수도권매립지의 정책결정과정에서 발생한 공공갈등 과정에서 정부정책에 반대하는 시민들의 참여 변화를 분석하였다. 시민참여의 변화는 일정한 시기를 두고 다양한 요인에 의해 역동적으로 변화하기 때문에 실증적으로 분석하는 데에는 한계가 있고, 기존 연구모형을 원용하여 분석하기에도 적절치 않아 근거이론을 적용하였다. 분석결과, 수도권매립지에 관한 갈등사례에서 시민참여는 ‘정부정책에 대항하지만 반대운동 세력이 미약하여 참여가 변화되는 특징’이 있음을 발견하였다. 정책결정은 권력이 강한 개인 혹은 집단의 의지가 반영되는데, 수도권매립지에 관한 정책결정은 민간 이해관계자들의 힘이 약하여 정부정책에 영향을 주지 못한 특징을 보였다. 한편, 이들은 서로 신뢰하는 수준이 낮았고 수도권매립지로 인한 보상제도에 문제가 있어 민민갈등이 발생하고 있었기 때문에 반대운동을 위한 광범위한 조직화를 이뤄내지 못하였다. 이러한 분석 결과를 통해서 본 연구는 공공갈등 구도에서 행위자들의 권력차이가 갈등 관계에 영향을 미칠 것이라는 가설을 제공하였다. 그리고 갈등관리에 필요한 절차적 정의를 갖추기 위해서 제도적으로 민간 이해관계자들에게 권한을 부여하는 시스템을 마련할 필요성을 제안하였다.

An improved empirical prediction method has been developed to predict the noise generated by impinging acoustic sources of flight vehicles. A representative empirical prediction method is the second distributed source method (DSM-II) in the report NASA SP-8072. However, since DSM-II is based on specific experimental data, it has limitations of not being able to perfectly reflect some acoustic phenomena such as impingement, reflection, and diffraction effects induced by interactive structures. In this study, we introduce some additional experimental data and a simple theory to the conventional method, and propose an improved empirical method that could reflect the above acoustic phenomena also. First, we describe the improved method that includes the impingement effect and validate the results for the Korean Sounding Rocket-III (KSR-III) in terms of overall sound pressure level (OASPL) and sound spectrum. Next, we expand this method to the case of a high speed vehicle surrounded by a canister that generates additional acoustic phenomena such as reflection and diffraction effects. We inspect the canister effects on the vehicle as the height increases, from the view point of OASPL and spectrum. Finally, the results are discussed.

Co-rotational formulation is one of the approaches applicable to geometrically nonlinear problems in structural analysis. Based on the assumptions of small degree of strain and large displacement, the CR formulation allows an accurate geometrically nonlinear structural analysis. In this paper, an improved CR formulation for dynamic nonlinear analysis is developed for the planar element. Also, extension of the local formulation is conducted for an elastoplastic problem. A simplified approach for planar elastoplastic analyses is proposed by maintaining the original feature of the CR formulation. Moreover, a modified approach in the local formulation is suggested in order to improve the level of accuracy in the plastic state. Contact analysis using the Lagrange multiplier is then implemented. During the present validation procedure of the proposed approaches, four different planar elements are employed. Those elements are investigated with regard to efficiency and accuracy in a general geometrically nonlinear and a geometrically nonlinear/elastoplastic problem, respectively. It is found that the present dynamic formulations gives both accuracy and consistency for all types of the plane element. Also, comparison regarding the choice of the local formulation for the elastoplastic problem is systematically conducted and demonstrated.

In this paper, structural modelling and dynamic analysis methods reflecting the characteristics of a liquid propellant were developed for a pogo analysis. The pogo phenomenon results from the complex interaction between the vehicle structural vibration in the longitudinal direction and the propulsion system. Thus, for an accurate vibration analysis of a liquid propellant launch vehicle, both the consumption of the liquid propellant and the change in the stiffness reflecting the nonlinear hydroelastic effect were simultaneously considered. A complete vehicle structure, including the liquid propellant tanks, was analytically modelled while focusing on pogo. In addition, a feasible liquid propellant tank modelling method was established to obtain an one-dimensional complete vehicle model. With these methods, comparative studies of the hydroelastic effect were conducted. Evaluations of the dynamic analysis of a reference vehicle were also conducted during the first burning stage. The numerical results obtained with the present orthotropic model and the dynamic analysis method were found to be in good agreement with the natural vibration characteristics according to previous analyses and experiments. Additionally, the reference vehicle showed the estimated occurrence of pogo in the first structural mode when compared with the frequencies of the propellant feeding system. In conclusion, the present structural modelling and modal analysis procedures can be effectively used to analyse dynamic characteristics of liquid propellant launch vehicles. These techniques are also capable of identifying the occurrence of pogo and providing design criteria related to pogo instability.

Various components of an engine nozzle are modeled as flexible multi-body components that are operated under high temperature and pressure. In this paper, in order to predict complex behavior of an engine nozzle, thermo-fluid-flexible multibody dynamics coupled analysis framework was developed. Temperature and pressure on the nozzle wall were obtained by the steady-state flow analysis for a two-dimensional nozzle. The pressure and temperature-dependent material properties were delivered to the flexible multi-body dynamics analysis. Then the deflection and strain distribution for a nozzle configuration was obtained. Heat conduction and thermal analyses were done using MSC.NASTRAN. The present framework was validated for a simple nozzle configuration by using a one-way coupled analysis. A two-way coupled analysis was also performed for the simple nozzle with an arbitrary joint clearance, and an asymmetric flow was observed. Finally, the total strain result for a realistic nozzle configuration was obtained using the one-way and two-way coupled analyses.

A nonlinear closed-loop load relief scheme is proposed to reduce the aerodynamic load during the ascent phase of a launch vehicle. The proposed controller is designed based on a back-stepping and sliding-mode control scheme with aerodynamic angle feedback. A hybrid load-relief strategy using the load relief scheme around the period of the maximum dynamic pressure and the traditional minimum-drift scheme during the other period is proposed. An aerodynamic angle estimator is also developed using a Kalman filter for the feedback of the load relief control. Numerical simulation is conducted to demonstrate the performance of the proposed strategy as well as the potential benefits.

Recently, there has been an increase in the research on flapping wing vehicles which mimic biological motions. One result has been the flapping wing micro-aerial vehicle. In this paper, the design requirements for flapping wing micro-aerial vehicles were established through an analysis with the unsteady blade element theory. Then, based on the flapping wing micro-aerial vehicle design requirements, a flapping wing mechanism using a pair of six-bar linkage was devised. Moreover, several candidates for the present mechanism were analyzed using a flexible multi-body dynamic analysis to ensure the structural appropriateness of the mechanism. By completing such procedures, the performance of the present mechanism could be evaluated. A detailed design was then conducted. The structural analysis of the present mechanism was conducted regarding its flapping operation in a vacuum. The resulting von Mises stresses in the linkage were targeted to be smaller than the yield stresses of the chosen material. Next, additional details of the design and an experiment on the present flapping wing micro-aerial vehicle were conducted to validate its performance.

High-speed flight vehicles (HSFVs) such as space launch vehicles and missiles undergo severe dynamic loads which are generated during the launch and in in-flight environments. A typical vehicle is composed of thin plate skin structures with high-performance electronic units sensitive to such vibratory loads. Such lightweight structures are then exposed to external dynamic loads which consist of random vibration, shock, and acoustic loads created under the operating environment. Three types of dynamic loads (acoustic loads, rocket motor self-induced excitation loads and aerodynamic fluctuating pressure loads) are considered as major components in this study. The estimation results are compared to the design specification (MIL-STD-810) to check the appropriateness. The objective of this paper is to study an estimation methodology which helps to establish design specification for the dynamic loads acting on both vehicle and electronic units at arbitrary locations inside the vehicle.

During the past few decades, various fluid–structure interaction (FSI) analysis approaches have been developed and applied to understand the physical processes related to a flexible flappin...

In this study, an equivalent structural modeling and structural analysis are proposed for the engine nozzle of a launch vehicle. A specific shell element that combines an optimal triangle (...

An insect-type flapping wing micro aerial vehicle offers high aerodynamic efficiency and maneuverability in confined spaces. The complicated aerodynamic/structural behavior of flapping wing micro aerial vehicle, however, causes difficulties regarding the dynamic control and parametric design. This paper develops a moderately accurate numerical framework taking into account the passive motion of the main wings. Finite-element-based multibody dynamics and two-dimensional unsteady aerodynamics are combined to simulate the hover of a flapping wing micro aerial vehicle. In addition, flexible and rigid wings are compared through numerical simulation considering the flexibility. In terms of the average thrust, numerical simulation by fluid–structure interaction shows good agreements against the experimental results within 5% discrepancy.

We present an all Eulerian approach to simulate the elastic response of a metal tube loaded explosively by a gaseous detonation. The high strain rate deformation of the metal tube subjected to high explosive detonation is mathematically described by hyperbolic processes where the characteristics of existing wave motions were correlated with the local particle velocities through the speed of sound in the metal. This is a favourable case for the hydrocode which is based on a compressible gas dynamics solver and for simulating a high strain rate and dominantly plastic response of a material subject to an explosive loading. The hydrocodes fall substantially short of predicting elastic motion without the plastic flow of the confining material, for relatively minor pressure loadings due to a gaseous explosion as opposed to a high explosive detonation of a charged tube. The corresponding loading pressure due to gaseous explosion is a few orders of magnitude lower than those resulting from high explosive loadings...

Recent advance in flapping-wing MAVs has led to greater attention being paid to the interaction between the structural dynamics of the wing and its aerodynamics, both of which are closely related to the performance of a flapping wing. In this paper, an improved computational framework to simulate a flapping wing is developed. This framework is established by coupling a preconditioned Navier–Stokes solution and a co-rotational beam analysis with a restrained warping degree of freedom. Validation of the present framework is performed by a comparison with examples from either earlier analyses or experiments. Further, a numerical analysis of a wing under simultaneous pitching and plunging motion is examined. The results are compared with those obtained with a wing under pure plunging motion, in order to assess the additional motion effect within a spanwise flexible wing. The comparison shows different aerodynamic characteristics induced by the flexibility of the wing, which can be beneficial.

In this paper, a triangular planar element is developed for a geometrically nonlinear structural analysis, which includes the drilling degrees of freedom using a corotational framework. Based on the assumptions of a small degree of strain and large displacement, the corotational framework allows an accurate geometrically nonlinear structural analysis. The presently improved corotational framework accommodates in-plane rotational behavior (that is, the drilling degrees of freedom) by using the corotational framework corresponding to a solidlike planar element. It focuses on triangular planar elements that will be useful for three-dimensional analysis using a reduced number of degrees while targeting a structure with a complex geometry, such as a flapping wing. Regarding the present analysis, validation by solving both static and time-transient problems is conducted. The fluid–structure interaction framework is then developed by using the present structural analysis. During this validation procedure, the pr...

본 연구에서는 정부에 대한 공정성 인식이 정부신뢰에 영향을 미치며, 이 과정에서 정부성과만족도가 매개효과를 갖는 것을 발견하였다. 먼저 정부의 공정성 변수를 절차, 분배, 상호작용 공정성으로 대별하여 정부신뢰 및 정부성과 만족도와의 관계를 살펴보았다. 그리고 사회를 신뢰하는 집단과 그렇지 않은 집단을 구분하여 집단 간 차이 분석을 실시하여 사회신뢰가 미약한 한국 사회에서 정부신뢰 향상을 위한 시사점을 도출하고자 하였다. 분석 결과, 정부신뢰에 가장 큰 영향을 미치는 변수는 상호작용공정성이었고, 정부성과만족도에 가장 큰 영향을 주고 있는 변수는 분배공정성이었다. 그리고 집단 간 차이분석에서는 독립변수가 절차공정성일 경우에 유의미한 집단 차이를 보여주고 있었다. 이러한 분석 결과를 근간으로 본 연구에서는 다음과 같은 시사점을 제시하였다. 첫째, 시민들에게 보여주는 공정하고 인간적 특징을 지닌 공감정부가 중요하며, 두 번째로 신뢰가 약한 한국 사회에서 정부가 적절히 운영되기 위해서는 절차적 공정성의 확보가 전제되어야 한다. 마지막으로 최근 논의 되고 있는 분배적 정의는 시민들에게 정부성과를 만족시킨 후에 정부신뢰를 향상시킬 수 있음을 실증적으로 증명하였다.

In this paper, the behaviour of a cantilevered beam was predicted to examine the difference between linear and non-linear static, dynamic analysis for a structure by using CR nonlinear formulation. Then, external transverse static and dynamic loads were applied at the free tip of the beam. Classical theories were used for the present linear analysis and co-rotational dynamic FEM program was used for the present nonlinear analysis. In the static analysis, effects of the load for the beam deflection were observed in both linear and nonlinear analysis. Then, normalized displacement at the tip of the beam was predicted for different frequency ratio and a significant difference was obtained in the vicinity of the resonant frequency. In addition, effects of frequency and time for the beam deflection were investigated to find the frequency delay. Keywords :linear analysis, nonlinear analysis, two-dimensional beam, natural frequency, harmonic excitation, co-rotational dynamic FEM, static analysis, frequency delay

There exist different kinds of aircrafts, such as conventional airplane, rotorcraft, fighter, and unmanned aerial vehicle. Their shape and feature are dependent upon their own assigned mission. One of the fundamental analyses performed during the aircraft design is the structural analysis. It becomes more complicated and requires severe computations because of the recent complex trends in aircraft structure. In order for efficiency in the structural analysis, a simplified approach, such as equivalent beam or plate model, is preferred. However, it is not clear which analysis will be appropriate to analyze the realistic configuration, such as an aircraft wing, i.e., between an equivalent beam and plate analysis. It is necessary to assess the limitation for both the one-dimensional beam analysis and the two-dimensional plate theory. Thus, in this paper, the static structural analysis results obtained by EDISON solvers were compared with the three-dimensional results obtained from MSC NASTRAN. Before that, EDISON program was verified by comparing the results with those from MSC NASTRAN program and other analytic solutions.

In this paper, Euler-Bernoulli beam theories(EB-beam) are used, and Fast Fourier Transformation(FFT) analysis is then employed to extract their natural frequencies using both analytical approach and Co-rotational plane beam(CR-beam) EDISON program. EB-beam is used to analyze a spring-mass system with a single degree of freedom. Sinusoidal force with various frequencies and constant magnitude are applied to tip of each beam. After the oscillatory tip response is observed in EB-beam, it decreases and finally converges to the so-called ‘steady-state.’ The decreasing rate of the tip deflection with respect to time is reduced when the forcing frequency is increased. Although the tip deflection is found to be independent of the excitation frequency, it turns out that time to reach the steady state response is dependent on the forcing frequency.

This study investigates the unsteady aerodynamic characteristics of the cambered wings of a flapping-wing micro air vehicle (FW-MAV) in hover. A three-dimensional fluid–structure interaction solver is developed for a realistic modeling of large-deforming wing structure and geometry. Cross-validation is conducted against the experimental results obtained also in the present study to establish more accurate analyses of cambered wings. An investigation is carried out on the unsteady vortex structures around the wings caused by the passive twisting motion. A parametric study is then conducted to evaluate the aerodynamic performance with respect to the camber angle at three different flapping frequencies including normal operating conditions. The camber angles producing the largest thrust and highest propulsive efficiency are estimated at each flapping frequency, and their effects on aerodynamic performance are analyzed in terms of the stroke phase. The timing and magnitude of the passive twisting motion, which are dependent on the camber angle at the operating frequency, greatly affects the unsteady vortex structure. Consequently, the camber angle designed at the operating frequency plays a key role in enhancing the aerodynamic performance of FW-MAVs.

This paper focused on improving a small-scaled helicopter rotor blade prototype with a flap-driving mechanism termed the Seoul National University Flap (SNUF). The design of the SNUF includ...

A methodology for determining the design allowables of composite laminates by using lamina level test data and finite element analysis (FEA) is proposed and verified in this paper. An existing method that yields the laminate design allowables by using the complete test results for laminates was improved to reduce the expensive and time-consuming tests. Input property samples for FEA were generated after considering the statistical distribution characteristics of lamina level test data. , and design allowables were derived from several FEA analyses of laminates. To apply and verify the proposed method, Hexcel 8552 IM7 test data were used. For both un-notched and open-hole laminate configurations, it was found that the design allowables obtained from the analysis correctly predicted the laminate test data within the confidence interval. The potential of the present simulation to substitute the laminate tests was demonstrated well.

This paper presents an all-direct domain decomposition approach for large-scale structural analysis. The proposed approach achieves computational robustness and efficiency by enforcing the compatibility of the displacement field across the subdomain boundaries via local Lagrange multipliers and augmented Lagrangian formulation (ALF). The proposed domain decomposition approach was compared to the existing FETI approach in terms of the computational time and memory usage. The parallel implementation of the proposed algorithm was described in detail. Finally, a preliminary validation was attempted for the proposed approach, and the numerical results of two- and three-dimensional problems were compared to those obtained through a dual-primal FETI approach. The results indicate an improvement in the performance as a result of the implementing the proposed approach.

In this paper, predictions of the tensile strength of carbon-fiber reinforced polymer (CFRP) laminate composites are attempted using an improved continuum damage mechanics model with emphasis on the material nonlinearity of the fiber, shear stiffness reductions, the damage contribution of matrix cracks in adjacent layers, and the statistical distribution of the fiber strength due to defects. The Weibull parameter to establish a continuum damage mechanics model was extended to five parameters to satisfy multiscale compatibility and interlamina effects. These five Weibull parameters, which serve as coefficients of a damage evolution function, were investigated by using a statistical model of progressive tensile fiber failure in a composite laminate. The Newton–Raphson method was utilized to formulate a damage-estimation method via a progressive failure analysis procedure. A continuum damage analysis based on the Matzenmiller–Lubliner–Taylor model was revised to add the statistical characteristics of the material strength. The proposed continuum damage mechanics method was also used to solve laminate tension problems with various stacking sequences. The verification test articles include 16- and 20-ply IM7/8552, AS4/8552, and T700/M015 carbon/epoxy composite laminate un-notched and open-hole tensile specimens. The strength prediction of each laminate specimen is presented.

A precise analysis model was developed in this paper to investigate the dynamic characteristics of rotating composite blades. An eighteen-node solid-shell finite element was used to model the blade structures. This study is focused on geometrically nonlinear problems, because the material is assumed linear elastic. Incremental total Lagrangian approach was adopted to allow estimations on arbitrarily large rotations and displacements. The equations of motion for the finite element model were derived by using Hamilton’s principle, and the resulting nonlinear equilibrium equations were solved by applying Newton-Raphson method combined with load control. A modified stress-strain relation was adopted to avoid the transverse shear locking problem, and fairly reliable results were obtained with no sign of locking phenomenon. The obtained numerical results were compared to several benchmark problems, and the results show a good correlation with the experimental data and other finite element analysis results. The vibration characteristics of shell- and beam-type blades were investigated. For shell-type blades, the dynamic characteristics may be significantly influenced by blade curvature, pre-twist, and geometric nonlinearity. For beam-type blades, one-dimensional beam and three-dimensional solid models offer comparable predictions for the straight and large aspect ratio blade. As blade aspect ratio decreases, considerable differences appear in the bending and torsion modes. The tip sweep angle tends to decrease the flap bending frequencies, but the torsion frequency increases with the tip sweep angle.

This paper presents validation of the helicopter rotor aeromechanics in descending flight. In its structural analysis, mixed variational geometrically exact beam model was used to capture nonlinear behavior of the rotor blade precisely. The present beam model was expressed by a mixed varational formulation in terms of the displacements, rotations, inertial forces, moments, and linear/angular momenta. Then displacements, inertial forces, and momenta can be directly extracted and solved simultaneously. And the finite-state dynamic inflow aerodynamics combined with blade element theory was used to predict induced inflow. The present results are compared by the wind tunnel test data for the higher-harmonic aeroacoustics rotor test (HART) II rotor with and without higher harmonic pitch control in descending flight. In the numerical results, trim control angles, section normal force, structural deflections, and structural loads were compared with those of the measurement. Predicted trim control angles and structural deflections show a good agreement with the measurement. But predicted flap bending moments show certain discrepancy with the measurement.

This paper presents a combined analysis for the rotor and dynamic components of a rotorcraft. The dynamic components consist of the following elements: a rotor shaft, a drive train, the engine shafts, and an engine and its governor. In this paper, the dynamic components were developed in a modularized fashion and combined with the rotor analysis. The present rotor structural analysis was derived based on the mixed variational formulation of moving beams and combined with finite-state dynamic inflow aerodynamics. A linear mass-spring-damper connection was used for a drive train representation and a simple linear control algorithm was used for an engine governor. To analyze the response of the components, a state-space equation was established based on the torque equilibrium. The present multi-component structural and aerodynamic analysis was solved for by measuring the steady-state trim and time-transient response. Numerical validation was performed using a sample hingeless rotor in a tiltrotor aircraft. The present prediction showed good agreement with CAMRAD II for both the trim analysis and the transient analysis with the trimmed state. And distinctions between the rotor with and without dynamic components were verified. Thus, an accurate framework for the rotor system including the dynamic components was developed, which would be useful during the preliminary design stage of creating a rotorcraft.