• Journal of Mechanical Science and Technology
• /
• 제16권3호
• /
• pp.327-337
• /
• 2002

This paper applies a recently developed on-line parameter identification (PID) technique to sets of real flight data and compares the results with those of a state-of-the-art off-line PID technique. The on-line PID technique takes Linear Regression from Fourier Transformed equations and the off-line PID is based on the traditional Maximum Likelihood method. Sets of flight data from the NASA F/A-18 High Alpha research Vehicle (HARV) circraft, which has been recorded from specifically designed maneuvers and used for our line parameter estimation, are used for this study. The emphasis is given on the accuracy and on-line measure of reliability of the estimates. The comparison is performed for both longitudinal and lateral-directional dynamics for maneuvers at angles of attack ranging u=20°through $\alpha$=40°. Results of the two estimation processes are also compared with baseline wind tunnel estimates whenever possible.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 한국추진공학회 2007년도 제29회 추계학술대회논문집
• /
• pp.335-338
• /
• 2007

This study carried out the 2D modeling for estimating the intake loss that is a important installed loss for the precise installed performance analysis of a gas turbine engine, and the 0D performance map that represents intake pressure loss change depending on flight Mach number and air mass flow rate was generated using the 2D modeling results. In order to evaluate the generation procedure of the intake performance map, the intake map generation was applied to a commercial aircraft intake configuration.

• International Journal of Control, Automation, and Systems
• /
• 제6권3호
• /
• pp.327-338
• /
• 2008

A fault tolerant control system design technique has been proposed and analyzed for managing performance degradation in the presence of multiple faults in actuators. The method is based on a control structure with a model reference reconfigurable control design in an inner loop and command input adjustment in an outer loop. The reduced dynamic performance requirements in the presence of different actuator faults are accounted for through different performance reduced (degraded) reference models. The degraded steady-state performances are governed by the reduced levels of command input. The reconfigurable controller is designed on-line automatically in an explicit model reference control framework so that the dynamics of the closed-loop system follow that of the performance reduced reference model under each fault condition. The reduced command input level is determined to prevent potential actuator saturation. The proposed method has been evaluated and analyzed using an aircraft example against actuator faults subject to constraints on the magnitude and slew-rate of actuators.

• International Journal of Aeronautical and Space Sciences
• /
• 제4권1호
• /
• pp.1-8
• /
• 2003

In this paper, a reconfigurable flight control system is designed by applying the sliding mode control scheme. The sliding mode control method is a nonlinear control method which has been widely used because of its merits such as robustness and flexibility. In the sliding mode controller design, the signum function is usually included, but it causes the undesirable chattering problem. The chattering phenomenon can be avoided by using the saturation function instead of signum function. However, the boundary layer of the sliding surface should be carefully treated because of the use of the saturation function. In contrast to the conventional approaches, the thickness of the boundary layer of our approach does not need to be small. The reachability to the boundary layer is guaranteed by the sliding mode controller. The fault detection and isolation process is operated based on a sliding mode observer. To evaluate the reconfiguration performance, a numerical simulation using six degree-of-freedom aircraft dynamics is performed.

• Journal of Aerospace System Engineering
• /
• 제8권1호
• /
• pp.30-35
• /
• 2014

This study has been conducted to develop an ejector system applied in the aircraft engine-bay ventilation system. Tandem-Ejector was selected as a component of ventilation system because it could achieve high ventilation performance in spite of motive flow with small flow rate. Tandem-Ejector is composed of a primary nozzle and two mixing ducts ($1^{st}$ mixing duct and $2^{nd}$ mixing duct). In this study, 1-D Tandem-Ejector model has been built with conservation laws and isentropic relation for 1-D ejector sizing and performance prediction. Computational Fluid Dynamics(CFD) has been conducted to investigate ejector performance and flow characteristics in the ejector. Also, Tandem-Ejector performance tests have been conducted to obtain ejector pumping performance and to investigate stand-off (gap between primary nozzle and $1^{st}$ mixing duct inlet) effect on ejector pumping performance.

• Journal of Institute of Control, Robotics and Systems
• /
• 제6권9호
• /
• pp.758-767
• /
• 2000

The properties of linear time-varying(LTV) systems vary because of the time-varying property of plant parameters. The generalized controller design method for linear time-varying systems does not exit because the analytic soultion of dynamic equation has not been found yet. Hence, to design a controller for LTV systems, the robust control methods for uncertain LTI systems which are the approximation of LTV systems have been generally ised omstead. However, these methods are not sufficient to reflect the fast dynamics of the original time-varying systems such as missiles and supersonic aircraft. In general, both the performance and the robustness of the control system which is designed with these are not satisfactory. In addition, since a better model will give the more robustness to the controlled system, a gain scheduling technique based on LTI controller design methods has been uesd to solve time problem. Therefore, we propose a new gain scheduled QFT method for LTV systems based on neural networks in this paper. The gain scheduled QFT involves gain dcheduling procedured which are the first trial for QFT and are well suited consideration of the properties of the existing QFT method. The proposed method is illustrated by a numerical example.

• Aerospace Engineering and Technology
• /
• 제5권2호
• /
• pp.240-247
• /
• 2006

Tilt rotor aircraft is a multi-configuration airplane which has three independent flight modes; helicopter, conversion, and aiplane. The control surface mixer resign is reqctired to generate and distribute efficient control forces and moments in each flight mode. In the conversion mode, the thrust vector is changed from helicopter mode to airplane, therefore the thrust vector makes undesired forces and moments which affect on pitch, roll and yaw dynamics. This paper describes the design results of control surface mixer design which minimize the undesired forces and moments due to nacelles tilting angle change for 4O% scaled model.

• Advances in aircraft and spacecraft science
• /
• 제3권2호
• /
• pp.149-170
• /
• 2016

Multi-Disciplinary Optimization (MDO) is widely used to handle the advanced design in several engineering applications. Such applications are commonly simulation-based, in order to capture the physics of the phenomena under study. This framework demands fast optimization algorithms as well as trustworthy numerical analyses, and a synergic integration between the two is required to obtain an efficient design process. In order to meet these needs, an adaptive Computational Fluid Dynamics (CFD) solver and a fast optimization algorithm have been developed and combined by the authors. The CFD solver is based on a high-order discontinuous Galerkin discretization while the optimization algorithm is a high-performance version of the Artificial Bee Colony method. In this work, they are used to address a typical aero-mechanical problem encountered in turbomachinery design. Interesting achievements in the considered test case are illustrated, highlighting the potential applicability of the proposed approach to other engineering problems.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• 제37권2호
• /
• pp.131-138
• /
• 2009

Analytical modal verification is considered as the process to provide an acceptable description of the subject structure's behaviour. In general, results of original analytical model are different with actual structure results to uncertainty like non-linearity of material, boundary and modified shape, etc. In this paper, the dynamic model of glider's wing is correlated with static deformation and vibration test results by goal-attainment method, multi-objects optimization technique. The structural responses are predicted by using finite element method and optimization is carried out by using the SQP(sequential quadratic programming) method which is widely used in the constrained nonlinear optimization problem. The MAC(Modal Assurance Criterion) is used to modify the mode shapes and quantify the similarity.

• Journal of the Korean Society for Precision Engineering
• /
• 제30권10호
• /
• pp.1061-1068
• /
• 2013

In this research, the design optimization of a composite sandwich has been performed for using as an airplane wing skin. Automated analysis framework for aero-structure interaction is used for calculating load data on the wing. For automated analysis framework, FLUENT is used for computational fluid dynamics (CFD) analysis. CFD mesh is generated automatically by using parametric modeling of CATIA and GAMBIT. A computational structure mechanics (CSM) mesh is generated automatically by the parametric method of the CATIA and visual basic script of NASTRAN-FX. The structure is analyzed by ABAQUS. Composite sandwich optimization is performed by NASTRAN SOL200. Design variables are thicknesses of the sandwich core and composite skin panel plies. The objective is to minimize the weight of the wing and constraints are applied for wing tip displacement, global failure index and local failure indexes.