• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.33 no.2
• /
• pp.98-105
• /
• 2005

This paper presents the result of the aeroelastic stability test of the small-scaled hingeless hub system with composite paddle blades in hover and forward flight conditions. Excitation tests of hingeless hub system installed in GSRTS(General Small-scale Rotor Test System) at KARI(Korea Aerospace Research Institute) were carried out to get lead-lag damping ratio of blades with flexures as hub flexure. MBA(Moving Block Analysis) technique was used for the estimation of lead-lag damping ratio. First, blades with metal flexures, then with composite flexures of the same dynamic properties of rotor system as metal one were tested. Tests were done on the ground and in the wind tunnel according to the test conditions of hover and forward flight, respectively. Composite flexures were found to have better damping characteristics over metal ones in the non-rotating vibration test, and it was confirmed that the use of composite flexures would give observable improvement in aeroelastic stability compared to metal ones in all test conditions.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.45 no.4
• /
• pp.269-275
• /
• 2017

The aeroelastic stability of a fighter type aircraft can be severly affected by the store mass, aerodynamic characteristics, and store combinations. Hence, the stability for the all store configurations must be substantiated before the aircraft in service. For the aeroelastic analysis, the design data and information for the aircraft structure, mass distribution, control surface characteristics, and external shape etc. are required. This is the reason that the store compatibility substantiations by a third party are restricted. However, according to the change of operational environment or the improvement of avionic technology, a new external store is developed and it should be installed on an aircraft without the support from the original supplier. This paper describe the process to substantiate the aeroelastic compatibility between a new external store and an imported aircraft whose design data is not available to a third party operating the aircraft.

• Journal of Aerospace System Engineering
• /
• v.14 no.5
• /
• pp.1-8
• /
• 2020

The aeroservoelastic analysis that deals with the interactions of the inertial, elastic, and aerodynamic forces and the influence of the control system have been performed. MSC Nastran was used for the free vibration analysis of the structure model as the pre-analysis. ZAERO was used to calculate the unsteady aerodynamic forces. The unsteady aerodynamic forces were verified by comparing with Doublet Hybrid Method. Karpel's Minimum-State Approximation method was used for approximation of the aerodynamic forces to the Laplace domain in the frequency domain. The aeroservoelastic state-space equation was obtained by combining the aeroelastic equation with the actuator dynamics. The analysis of aeroservoelastic stability concerning the elevator input of the high aspect ratio model was performed. The root-locus method and time-integration method were used for the analysis of aeroservoelastic in frequency and time domain.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.35 no.6
• /
• pp.503-508
• /
• 2007

In this study, the aeroelastic response and stability of bearingless rotors are investigated using a large deflection beam theory. The outboard main blade, flexbeam, and torque tube are all assumed to be an elastic beam undergoing arbitrary large displacements and rotations. The finite element equations of motion obtained from Hamilton's principle. Two-dimensional quasi-steady strip theory is used to evaluate aerodynamic forces. In hover, the modal approach method based on coupled rotating natural modes is used for the stability analysis. In forward flight, the nonlinear periodic blade steady response is obtained by integrating the full finite element equation in time through a coupled trim procedure with a vehicle trim. The results of the full finite element analysis using the large deflection beam theory are compared with those of a previously published modal analysis using the moderate deflection-type beam theory.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.43 no.9
• /
• pp.781-786
• /
• 2015

An aeroelastic stability can be degraded due to an aeroelastic modeling error and a structural uncertainty. Therefore it is necessary to predict the aeroelastic stability boundary considering an aeroelastic modeling error and a structural uncertainty. Robust aeroelastic analysis was proposed to predict the aeroelastic stability boundary considering these error and uncertainty. In the present study, the robust aeroelastic modeling and analysis were performed by using the ${\mu}$ analysis technique and the aeroelastic model of the control fin with modal approach and MSA. The computer program for the robust aeroelastic analysis was developed and verified by comparing its results with those of conventional aeroelastic analysis methods.

• Composites Research
• /
• v.17 no.3
• /
• pp.29-37
• /
• 2004

The aeroelastic stability analysis of composite bearingless rotors is investigated using a large deflection beam theory in hover. The bearingless rotor configuration consists of a single flexbeam with a wrap-around type torque tube and the pitch links located at the leading edge and trailing edge of the torque tube. The outboard main blade, flexbeam and torque tube are all assumed to be an elastic beam undergoing flap bending, lead-lag bending, elastic twist and axial deflections, which are discretized into beam finite elements. For the analysis of composite bearingless rotors, flexbeam is assumed to be a rectangular section made of laminate. Two-dimensional quasi-steady strip theory is used for aerodynamic computation. The finite element equations of motion for beams are obtained from Hamilton's principle. The p-k method is used to determine aeroelastic stability boundary. Numerical results are presented for selected bearingless rotor configurations based on the lay-up of laminae in the flexbeam and pitch links location. A systematic study is made to identify the importance of the stiffness coupling terms on aeroelastic stability for various fiber orientation and for different configuration.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.47 no.10
• /
• pp.695-704
• /
• 2019

In hypersonic regime, the complicated interaction between the air and surface of aircraft results in intensive aerodynamic heating on body. Provided this phenomenon occurs on a hypersonic vehicle, the temperature of the body extremely increases. And consequently, thermal deformation is produced and material properties are degraded. Furthermore, those affect both the aerothermoelastic stability and thermal safety of structures significantly. With the background, thermal safety and dynamic stability are studied according to the altitude, flight time and Mach number. Based on the investigation, design guideline is suggested to guarantees the structural integrity of hypersonic vehicles in terms of both of thermal safety and dynamic stability.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.39 no.4
• /
• pp.297-305
• /
• 2011

In this study, the aeroelastic analysis of rotorcraft in forward flight has been performed using dynamic inflow model to handle unsteady aerodynamics. The quasi-steady airload model based on the blade element method has been coupled with dynamic inflow model developed by Peters and He. The nonlinear steady response to periodic motion is obtained by integrating the full finite element equation in time through a coupled trim procedure with a vehicle trim for stability analysis. The aerodynamic and structural characteristics of dynamic inflow model are validated against other numerical analysis results by comparing induced inflow and blade tip deflections(flap, lag). In order to validate aeroelastic stability of dynamic inflow model, lag damping are also compared with those of linear inflow model.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.36 no.2
• /
• pp.163-170
• /
• 2008

The aeroelastic stability analysis of a soft-in-plane, composite hingeless rotor blade in hover and in forward flight has been performed by combining the mixed beam method and the aeroelastic analysis system that is based on a moderate deflection beam approach. The aerodynamic forces and moments acting on the blade are obtained using the Leishman-Beddoes unsteady aerodynamic model. Hamilton's principle is used to derive the governing equations of composite helicopter blades undergoing extension, lag and flap bending, and torsion deflections. The influence of key structural modeling issues on the aeroelastic stability behavior of helicopter blades is studied. The issues include the shell wall thickness, elastic couplings and the correct treatment of constitutive assumptions in the section wall of the blade. It is found that the structural modeling effects are largely dependent on the layup geometries adopted in the section of the blade and these affect on the stability behavior in a large scale.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
• /
• 2006.05a
• /
• pp.680-685
• /
• 2006

This paper describes the aeroelastic stability test of the small-scaled 'Next-Generation Blade(NRSB)' with NRSH (Next-Generation Hub System) and HCTH hingeless hub system in hover and forward flight conditions. Excitation tests of rotor system installed in GSRTS(General Small-scale Rotor Test System) at KARI(Korea Aerospace Research Institute) were tarried out to get lead-lag damping ratio of blades with flexures as hub flexure. MBA(Moving Block Analysis) technique was used for the estimation of lead-lag damping ratio. First, NRSB-1F blades with HCTH hub system, Then NRSB-1F with NRSH hub system were tested. Second, NRSB-2F blades with NRSH hub system were tested. Tests were done on the ground and in the wind tunnel according to the test conditions of hover and forward flight, respectively. Non-rotating natural frequencies, non-rotating damping ratios and rotating natural frequencies were showed similar level fir each cases. Estimated damping ratios of NRSB-1F, NRSB-2F with HCTH and NRSH were above 0.5%, and damping ratio increased by collective pitch angle increasement. Furthermore damping ratios of NRSB-2F were higher than damping ratios of NRSB-1F in high pitch angle. It was confirmed that the blade design for noise reduction would give observable improvement in aeroelastic stability compared to paddle blade and NRSB-1F design.