© 2023, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.Operational modal analysis (OMA) has been widely used in many fields of study because it allows identifying the modal parameters of a flexible structure in its operating condition. The system is under unknown working loads assumed to be random with broadband spectral characteristics. These hypotheses are not always easy to fulfill, generating uncertainty about identified modal parameters. This study evaluates and compares the effectiveness of two OMA techniques, enhanced frequency-domain decomposition (EFDD) and Ibrahim time domain (ITD), in the accuracy of modal parameter estimation of an unmanned aerial vehicle (UAV) structure with output-only data obtained by flight testing. To evaluate the influence of the number of sensors used in the identification of the modes, different measurements setups were considered to carry out in-flight modal identification analyses. Some works have addressed uncertainty by focusing on retesting or subdivision of a single measurement record. This work innovates in presenting an uncertainty study considering the variables that intervene in the estimation of PSD. The uncertainty in the identified modal parameters is obtained using the variability of the values of the parameters found. The modal frequencies values observed employing EFDD and ITD do not present substantial variations associated with the PSD matrix estimates. The EFDD damping ratio values show significant variability because they are mainly affected by spectral leakage, while the ITD damping ratio values are less sensitive to Welch’s method parameters variation. The root mean square deviations (RMSDs) of the frequencies values for both techniques are compared with those resulting from ground vibration testing.

© 2023, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Synthetic Air Data Systems are air data parameters real-time estimation algorithms. Estimation of such parameters have been under study for a few decades. System Identification theory gives some tools for both time and frequency domain. Several studies have been conducted to investigate this problem in the time domain, but the applicability of frequency-domainal gorithms is still to be investigated. This work proposes a frequency-domain formulation for the synthetic air data problem, which is validated using a time-domain method (Recursive LeastS quares). Both methods are applied to real flight test data and estimation results are discussed. Effects of the availability of side-slip parameter are evaluated and estimates uncertainties due to model parameters accuracy (stability derivatives) are also presented.

Copyright © (2022) by International Council of Aeronautical Sciences (ICAS) All rights reserved.There are some approaches for updating models to later model the aeroelastic behavior, and in this work, the Modal Assurance Criterion (MAC) helps identify the parameters. The objective of this work was to update the finite element model for the EOLO aircraft. We used the modal shapes derived from Ground Test Vibration (GVT) as a basis of comparison for the MAC, in addition to using the Nastran software to optimize the stiffness properties of the analytical model of the EOLO aircraft. It noted that the natural frequencies of the updated model approached the GVT data and the cross-correlation improved, but the correlation was far from ideal. Therefore, the model was updated and improved over the initial model.

© (2022) by International Council of Aeronautical Sciences (ICAS) All rights reserved.This paper presents the identification of an Unmanned Aerial System (UAS) with flexible wings from open-loop data using subspace methods. For aerodynamic and flight control systems, a reliable model is important to comprehend the system behaviour and to design a feedback loop, that can be applied as well to minimize the effects of structural flexibility. So, a parametric model identification for flexible aircraft applying subspace techniques was performed. Preliminary results presented in this paper are related to identification using synthetic data. Finally, it is shown the experimental results from the first flight test performed in open-loop operation. The experimental results reveals that subspace methods estimate a state-space model suitable, with better fit for the range of frequencies of the experimental data. Therefore, it was not possible to obtain a representative model for a broader frequency range, however, this is not a limitation of the method, but a persistent excitation problem associated with the limited frequencies in the input signal.

© (2022) by International Council of Aeronautical Sciences (ICAS) All rights reserved.Synthetic Air Data Systems are airspeed estimation algorithms. Such algorithms are built using measurements from sensors other than the classical Pitot tubes, from which airspeed estimates can be computed. This paper presents brief discussions over three direct estimation algorithms that use inertial sensors (IRS) and GPS as sources of information. It is also proposed and tested a recursive airspeed and thrust estimation (RATE) algorithm. Finally, a simple implementation using Extended Kalman Filter (EKF) is tested and results are compared. The possibility to use angle-of-attack and/or temperature probes is also discussed. We also discuss practical aspects regarding airspeed, altitude and temperature estimations.

© (2022) by International Council of Aeronautical Sciences (ICAS) All rights reserved.Subscale aircraft have been used for decades to design new aircraft and evaluate new design techniques. The acquisition of in-flight data from subscale aircraft is already possible today, such as a manned or fullscale aircraft. Thus, more reliable flight simulators are built for flight quality analysis and control design. This work aims to implement a data acquisition and processing system, with the objective of identifying the complete dynamics of a subscale aircraft, model Cessna 182.

© 2022, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.The development of microelectronics combined with the cheapness of manufacturing processes has allowed the construction of subscale models equipped with sensors and control systems equivalent to a real aircraft. This work analyses the Generic Future Fighter (GFF) subscale concept developed by Linköping University under the Future Aircraft Design and Demonstration (FADEMO) project. The GFF subscale is a radio-controlled aircraft with 14% of the size of the full-scale concept aircraft. A Stability Augmentation System (SAS) will be designed to stabilize the longitudinal dynamics for different positions of the c.g., artificially modified for three different positions. Despite the several control techniques currently available, methods such as the Linear Quadratic Regulator (LQR) are still adopted for the stability control of aircraft in flight. However the LQR method present in their classic form, limitations to incorporate performance parameters and operational restrictions in the design phase. A promising alternative to circumvent this problem is the use of Linear Matrix Inequalities (LMIs) as a tool to convert stability and control problems into optimization problems. This work presented an LQR-LMI formulation augmented by D-stability criterion to simplify the determination of a single feedback gain matrix that guarantees the stability and keeps the flight characteristics by varying the c.g. position.

© 2021 32nd Congress of the International Council of the Aeronautical Sciences, ICAS 2021. All rights reserved.With an always increasing demand for more efficient aircraft due to both economic and environmental purposes, academy and industry are studying hybrid-electric and full-electric concepts to explore new aircraft design opportunities. This paper expands on the results of previous publications, using a Cessna 208B Grand Caravan as a platform for the implementation of distributed electric propulsion to enable the use of high-lift propellers. The design space is swept to evaluate the impacts of the technology in wing weight, propulsive system sizing and weight as well as in payload, range and fuel consumption on different simulated missions. The models are integrated using SUAVE, a conceptual level design environment, which is used to integrate the aircraft model and run the simulations. Results show that generating enough extra lift with the propellers require large amounts of power, resulting in a heavy propulsive system that hinders the payload and range capabilities of the aircraft and are unfortunately not compensated by the small aerodynamic gains generated by this configuration.

© 2021 32nd Congress of the International Council of the Aeronautical Sciences, ICAS 2021. All rights reserved.In the development of automatic piloting systems nowadays, flight tests are required to validate operations and tuning of the control loops, making the process costly. For an optimal point of stability within a region of operation, this work aims to use the modern control technique of Linear Quadratic Regulator (LQR), with minimization through the Riccati equation for the optimization of Proportional, Integral, and Derivative (PID) control loops in longitudinal piloting. The aircraft considered for the flight tests was the C2 fixed-wing Unmanned Aerial Vehicles (UAV) used for agricultural purposes. The nonlinear model coefficients of the aircraft were acquired using well-known computational methods. The inertia properties were acquired through drawings made in Computer-Aided Design (CAD) with the Catia® software and the aerodynamic properties' estimation with the Omni3d® software. The aircraft's applied system was the Micropilot® LRC2 autopilot that has cascade PID control loops for altitude and trajectory control; however, the control loops tuning responsible for longitudinal movement are this work's main contribution. In parallel, a flight test campaign was carried out to collect data and tune the autopilot gains in flight by an empirical method based on Ziegler-Nichols' method. The gain data collected during the flights are used to compare the data obtained by the theoretical computer model of the aircraft. Different performances related to the gains obtained by the flight test and the PID control loops' optimization method through the LQR method are demonstrated with the non-linear model's application under the effects of disturbances. The main achieved results are about the minimum energy cost. This minimization in energy cost is due to the PID in-flight tuning that takes more energy in the actuators than the PID optimized with the LQR method, which proves to be a promising system for faster development of agricultural UAVs.

© 2021, Journal of Aerospace Technology and Management. All rights reserved.Combustion instability can severely impair the operation of many kinds of combustion engines. Acoustic resonators are widely used to suppress the pressure oscillations caused by the coupling between the combustion process and the combustion chamber acoustic modes. Combustion chambers with subsonic flow in its inlets and outlets, like gas turbine combustors, exhibit some acoustical damping due to the presence of openings. In such chambers, the acoustic modes are complex. In a complex mode, the antinode regions can be shifted from its position in the corresponding real mode. In this work an experimental acoustic modal analysis of a cavity with an opening was performed. Acoustic frequency response functions were obtained by using a volume acceleration source, a microphone and a data acquisition system. The PolyMAX algorithm was used to estimate longitudinal modes in its real and complex versions. A comparison was performed and the results show that, for some modes, the antinode region placement could change reasonably. This suggests that the use of complex modes for location of antinode regions provides more accurate results and consequently could be a better way to identify positions, where resonators provide maximum damping in order to minimize combustion instability in subsonic combustion chambers.

© 2021, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Operational Modal Analysis (OMA) is a methodology to obtain the modal properties of a structure using the dynamical responses of the system only. That methodology is very useful in aeroelastic in-fligh testing where there are difficulties in measuring directly the aerodynamic loads on aircraft. For risk reduction in the future flight operations of the EOLO at ITA, wind-tunnel tests were performed. EOLO is an Unmanned Aerial Vehicle (UAV) with high aspect ratio and structural flexibility, designed to study aeroelastic phenomena and to evaluate the interaction of flexible effects with the aircraft flight dynamics. In this work The Frequency Domain OMA techniques: Decomposition (FDD), Enhanced Frequency Domain Decomposition (EFDD), Spatial and Frequency Domain Decomposition (SFDD) are applied to the vibrational data from wing tunnel test campaign in different operation conditions, using acceleration and strain measurements. The results between the OMA techniques are compared with previous modal characterization from Ground Vibration Test (GVT) and numerical aeroelastic analysis. The different generations of the frequency domain decomposition methods proved their suitability for use in aircraft aeroelastic characterization using accelerations and strain measurements.

© 2020 Elsevier LtdThis paper shows the results of traditional aerodynamic analysis and the flexible flight dynamics modeling including the effects of structural motion for a flexible wing unmanned aerial vehicle. The influence of some design parameters such as wing flexibility, horizontal/vertical tail aerodynamics is investigated for aeroelasticity and flight dynamics of flexible aircraft. The research platform is an Unmanned Aerial Vehicle (UAV), made of composite material. Its wing span is 4 m and reaches a high aspect ratio, whose value is 18.9. For a traditional analysis, the Vortex Lattice Method (VLM) was used to obtain conventional aerodynamic and control derivatives. The flexible flight dynamics model is based on the work of Waszack and Schmidt. In this approach, it is used the mean-axes reference system and it is assumed that structural deformations is small and described by a set of eigenmodes. The dynamics model incorporates the first normal modes obtained by Ground Vibration Test (GVT) campaigns. A focus of the paper lies on providing a useful model for dynamics system identification and in-flight aeroelastic testing. A developed platform for in-flight system identification and the acquisition system is described. A flight path reconstruction process from a flight test campaign results is shown.

© 2020 AIAA.With an always increasing demand for more efficient aircraft due to both economic and environmental purposes, academy and industry are studying hybrid-electric and full-electric concepts to explore new aircraft design opportunities. This paper proposes a study based on a Cessna 208B Grand Caravan, using it as a platform to implement distributed electric propulsion and enable the use of high-lift propellers by electrifying the propulsive system. Key design parameters of the aircraft are varied to evaluate the effectiveness of the lift augmentation system as well as its effects on generated thrust and aerodynamic efficiency. The effects of the propellers slipstreams on the wing are implemented on SUAVE, a conceptual level design environment, which is used to integrate the aircraft model and run the simulations. Results of the analyses differ from what is available on the literature, yielding aerodynamic efficiency gains that are much more modest than what was expected according to assumptions made on recent publications.

© 2020 Proceedings of ISMA 2020 - International Conference on Noise and Vibration Engineering and USD 2020 - International Conference on Uncertainty in Structural Dynamics. All rights reserved.A Bond Graph model for a light business aircraft and its brake system has been developed. The dynamic response of the aircraft is analyzed with two degrees of freedom considering it as a rigid body. Dual stage servo-valves, the hydraulic tubing and hoses, and brake assemblies behavior are considered on the analysis. The simulations are run using 20-Sim® software. Braking performance efficiency is assessed considering three different methods: Torque Method (AC 25-7D), Stopping Distance Efficiency and Developed μ Efficiency Methods (SAE AIR1739B). For each method, an On-Off pressure and slip control technique is implemented. The analysis of the results provides key insights for a better correlation between aircraft braking performance top level requirements and brake system specification. Therefore, this paper supports system and performance engineers on the definition of more accurate requirements on the aircraft early design phases.

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.With an always increasing demand for more efficient aircraft due to both economic and environmental purposes, academy and industry are studying hybrid-electric and full-electric concepts to explore new aircraft design opportunities. This paper proposes a study based on a Cessna 208B Grand Caravan, using it as a platform to implement distributed electric propulsion and enable the use of high-lift propellers by electrifying the propulsive system. Key design parameters of the aircraft are varied to evaluate the effectiveness of the lift augmentation system as well as its effects on generated thrust and aerodynamic efficiency. The effects of the propellers slipstreams on the wing are implemented on SUAVE, a conceptual level design environment, which is used to integrate the aircraft model and run the simulations. Results of the analyses differ from what is available on the literature, yielding aerodynamic efficiency gains that are much more modest than what was expected according to assumptions made on recent publications.

© 2019 IEEE.The present paper proposes the application of Model Predictive Control (MPC) to the bipedal walking problem. Classically, bipedal robots maintains constant height of the center of mass (CoM) during walking, since this constraint makes the underlying dynamical system linear. Nevertheless, researches show that vertical CoM motion is one of many mechanisms humans use to reduce energetic cost during walking. In this paper, we show that if the height is modified through a predefined function, the system becomes linear time-varying, which may be handled by MPC techniques. By means of simulations, the stability behavior of the robot is verified. Finally, a high-fidelity simulation model based on the Gazebo simulator is used to validate the energetic cost reduction due to the vertical CoM motion.

© 2019 American Institute of Aeronautics and Astronautics. All rights reserved.This paper presents an aerodynamic analysis of an airship through a computational tool, aiming to obtaining aerodynamic coefficients to be used in the design of novel airship models. A comparison between the model from simulation and from the wind tunnel test is used to validate the computational method. Due to the geometry presented by airships the analytical estimation of their coefficients becomes extremely difficult. Thus, computational tools can assist in the preliminary design of this type of air vehicle, as well as providing preliminar aerodynamic parameters, reducing the number of batteries required for wind tunnel tests. The description of the aerodynamic behavior of the airship was performed using the XFLR5 software, the modeling and setup details used will be presented in the course of this paper. The results obtained were validated through the comparison analysis presented between simulations and previous experimental results obtained in a wind tunnel using the YEZ-2A airship.

© 2019 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.This paper presents a methodology employed to characterize the aeroelastic behavior of an Unmanned Aerial Vehicle (UAV) with high aspect ratio and structural flexibility. Commonly, modal characteristics of aircraft are experimentally obtained by means of a Ground Vibration Test (GVT), which in turn could be used to correlate and update numerical models and further applied to aeroelastically characterize the aircraft. The input-Output experimental modal analysis is not easily applied to aeroelastic tests in actual flight conditions because of the difficulties measuring the actual inputs due to aerodynamic loads. Therefore, the Ouput-Only (O-O) approach also known as Operational Modal Analysis (OMA) was used to identify the modal characteristic of aircraft using accelerometers and strain sensors. In this study the OMA based on the Enhanced Frequency Domain Decomposition (EFDD) was applied to process the operational vibration data. This paper discusses and compares the results between Input-Output modal analysis and OMA approaches. A numerical flutter analysis was performed to observe the evolution of aeroelastic damping and frequencies as a function of airspeed, and to understand coupling mechanisms. The Aircraft aeroelastic behavior is studied for different flight conditions. The g-method for aeroelastic stability analysis was employed. Based on information from the GVT a flight test planning was conducted. The data acquisition system is described below.

© 2019 American Institute of Aeronautics and Astronautics. All rights reserved.This paper proposes a methodology based on multidisciplinary optimization for sizing the fins of a Lighter-Than-Air (LTA) vehicle that merges the airship and tethered aerostat functionalities. The purpose of this design is to improve the airship’s stability in a tethered flight condition by expanding fin dimensions. Therefore, an optimum sizing of the fins that meets the minimum stability requirements of both free and tethered flight conditions is approached. This study integrates in a concurrent way the mathematical modeling of the tethered airship and the methodology for fins sizing. A design vector containing initial geometric parameters of the airship with different fins size is optimized with respect to the aerodynamic stability of the tethered airship. The parameters related to the empennage is varied restricted to parameterized equations that describe the airship dynamics in free and tethered flight condition. The stop criterion of the optimization processes is obtained when the airship dynamics meets the minimum requirements used to compose the cost function based on the aerodynamic mode damping, resulting in the optimized dimension of the airship fins. This design enables to expand the airship flight envelope, ensuring a stable flight when hovering in a tethered flight condition and ensuring maneuverability to the airship for performing a free flight.

© 2017 IEEE.A methodology was developed in this work for the automatic counting of individual seedlings in plantations of Eucalyptus spp from high definition photographs with the help of Scientific Python Libraries from literature. The problem to be investigated was presented and two different ways of solving it were discussed together with their implications. With the algorithm properly validated on training data, an actual business case of seedlings detection and counting out of a mosaic aerial image was proposed as testing data. The high-definition pictures were taken by multispectral sensor onboard an UAV from an Eucalyptus spp plantation stand of approximately 25 hectares and provided by Eldorado Brasil. The results were considered very encouraging, stimulating future works in this line of research.

© 2018, The Brazilian Society of Mechanical Sciences and Engineering.Due to the increase in aircraft systems complexity along the decades and the continuous certification requirement improvements for safer operations, the safety assessment accomplished by systems engineers has been demanding more effort from the specialists to make a complete evaluation of the system and respective interfaces. The capability of predicting the real effects of components failures in the system behavior to make better assessments of their severities and to support troubleshooting processes during aircraft operation has also represented a challenging activity. In that context, the development of computational models and simulation has become a common practice in the industry. Therefore, the aim of the present work was to demonstrate the benefits of working in a cohesive manner with two particular modeling techniques: a physical modeling based computational software and the bond graph concepts, to enhance the specialist’s comprehension about the impacts of particular failures in system performance. As a case study, an aircraft hydraulic brake system has been chosen since it performs important, safety-related functions in aircraft operation. For that purpose, a computational model parameterized in LMS Amesim® software is used, after a deep validation process, to assess the behavior of system relevant variables in normal and faulty operating conditions. In parallel, a bond graph diagram representative of a system component is applied as a support tool to assess typical failure modes and help selection of relevant ones for simulation.

© 2018 IEEE.Shape memory alloy (SMA) wires have extensive use in many areas of the industry nowadays and its development continues reaching new applications as studies progress. This paper proposes a SMA measurement device that uses affordable components, such as the Arduino micro-controller and a LabVIEW programming language interface. With an antagonistic mechanism design, data on temperature, strain and stress is acquired to confirm the measuring capabilities of the full equipped instrument, rendering visualizations of phase transformations and opening way for further development in control and detailed acquisition of shape memory alloy wire properties.

© 2018, The Brazilian Society of Mechanical Sciences and Engineering.In this work, the authors propose a formation control strategy of a group of three multirotor aerial vehicles being able to avoid multiple obstacles and collisions. To deal with this problem, a decentralized architecture is proposed which has one model predictive controller per vehicle including a set of convex constraints on the vehicle’s position to prevent collisions with other agents and different shapes of obstacles. The resulting decentralized scheme controls the formation based on a virtual structure approach. For the purpose of avoiding collisions, each local controller considers the predicted position of every neighbor vehicles. The effectiveness of the developed scheme is demonstrated through numerical simulations considering a “figure-of-eight” as the reference trajectory, and the results show its capability to handle thrust force, obstacle and collision avoidance constraints.

© 2017, Springer Science+Business Media B.V.This paper presents position control of the control surface of an aeroelastic typical section using shape memory alloy (SMA) actuators. The actuator was designed using an antagonistic pair of SMA wires, thus allowing its use to move the control surface to up and down. The mathematical model adopted here describes the relations among aeroelastic section, models of SMA wire heat convection, constitutive law and phase transformations of the SMA. Three types of controllers are presented based on variable structure control approach and the deflection of the control surface via two pairs of antagonistic wires are made in a closed loop system. The effectiveness of the actuation system in positioning the control surface is evaluated numerically.

© 2003-2012 IEEE.The present work describe the use of different maneuvers with intention to improve the identification of latero directional model of an Unmanned Aerial Vehicle (UAV). The best flight datas was selected according to criteria described in this paper. It was used the methodology of 4 M's (Maneuver, Model, Method and Measure) identification, to estimate the parameters of the model.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Morphing wings can optimize their performance during the whole mission profile. Using a morphing wing can make landing speeds lower and flying safer, make wings produce less noise and drag and reduce fuel consumption. This paper describes how a morphing wing is designed, constructed and tested. Table tests show a significant change in air profile is achieved within 1 second of actuation. Wind tunnel tests show that morphing the wing shifts the Cl − α graph. This means that morphing the wing performs the same function as actuating a flap. Therefore, using SMA wire to create a morphing wing is possible and it is proven that morphing wings provide benefits over a normal wing.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.This work presents a methodology for system identification of a large flexible aircraft operating in closed-loop. The feedback in some cases is necessary because the system in open-loop is unstable or because a controller, known or not, is present on system and is not possible to removed it. The synthetic data of a nonlinear dynamic to the aircraft considering three symmetric and two anti-symmetric flexible modes are used in the identification algorithm. The identification algorithm is a non iterative subspace method well applied for both open and closed-loop data. The preliminary results suggest a representative model for the aircraft, obtaining the state-space matrices that are of very interest and used for control system analysis and design.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.This paper proposes a novel LTA platform, the tethered airship that is a merge of airship and tethered aerostat functionalities, combining the navigation abilities of airships and the hovering performance of tethered aerostats that is designed to remain stationary in high wind conditions, to transfer data with high speed data link to the ground base through an electric tether and to fly unlimited time periods. This paper presents the sizing based in a scaled model of the YEZ-2A airship, the designing and fabrication of each component, and the integration that originates the novel tethered airship prototype in which it is validated through flight experiments.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.The present work describes the system identification process of the lateral-directional stability derivatives of an Unmanned Aerial System (UAS) [1]. The Maneuver, Model, Method, Measures and Validation (M4V) [2], is a well known in-flight identification methodology that was applied to the VECTOR-P UAS. The maneuvers adopted to excite the lateral modes of the system were evaluated with the energy spectral density (ESD). The data was acquired during flight tests by the data acquisition system specifically developed to the Vector-P. Finally, the validation of the identified parameters was performed using statistical methods.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.In the context of system failures analysis, the present work aims to demonstrate the benefits of working in a cohesive manner with two particular modeling techniques, a physical modeling-based computational software and the bond graph concepts, when identifying failure modes and assessing the impacts of typical failures in an aircraft hydraulic brake system. The brake system performs an important, safety-related function in aircraft operation.

© Proceedings of ISMA 2018 - International Conference on Noise and Vibration Engineering and USD 2018 - International Conference on Uncertainty in Structural Dynamics. All rights reserved.Due to the increasing complexity of aeronautical systems, it became more and more important to detect possible failures, avoiding costs with maintenance and time out of operation. Modeling techniques and computational softwares made possible to analyze systems behaviour under normal and failure conditions, helping to prevent these problems. In this work, an aircraft anti-skid brake system is considered as a study case. Therefore, the aircraft brake dynamics and the brake system are modeled using Simulink. In the brake model, some common faults are introduced in order to observe its impacts on the braking performance. A model based fault detection and isolation (FDI) method using analytical redundancy relations (ARRs) is proposed. ARRs are equations relating the system constraints. The numerical evaluation of these equations generates residuals indicating the system deviation from its normal operation. The coupling of the Simulink behavior model with the ARRs is presented, permitting the residuals analyses for each failure mode.

© Springer International Publishing AG, part of Springer Nature 2019.This work describes the nonlinear identification applied to an aeroelastic pitch-plunge system using polynomial NARMAX model and a stability analysis. The apparatus is available and consists of a wing typical section with pitch and plunge degrees of freedom. The identification procedure aims to obtain the parameters for the mathematical model including the torsional stiffness as a quadratic polynomial function. The candidate structure to the polynomial model is obtained from discretization of a continuous-time state-space model and the predictions are obtained via the identification procedure using simulated data. The simulation is performed considering the aerodynamics with free stream velocity increased within an established velocity range which includes the flutter phenomenon. In future work, a data acquisition from the experimental apparatus will be performed. The NAR-MAX model indicates a polynomial function of fourth order for the nonlinearity and a stability analysis, discussed in this work, mapping the nonlinear regions.

© 2017 IEEE.Due to low-cost and simplicity, differential drive mobile robot are very popular in academic and hobby environments. However, in some applications, such as robot soccer, these robots need to move quickly and aggressively, thus requiring the use of techniques from control systems theory, where an accurate mathematical model is paramount. On the other hand, given budget limitations, the components used for academic robotics competitions are often acquired from hobby-grade manufacturers, which usually provide incomplete specifications. In this paper, we introduce a procedure to experimentally measure the parameters involved in the dynamical model of a differential drive robot. The procedure is intentionally based on a low-cost setup, which uses equipments found in academic laboratories. Moreover, we show experiments results that validate the mathematical model obtained through the experimental procedure.

© 2017 IAAThe primary purpose of this paper is to provide insight into control-structure interaction for satellites comprising flexible appendages and internal moving components. The physical model considered herein aiming to attend such purpose is a rigid-flexible satellite consisting of a rigid platform containing two rotating flexible solar panels. The solar panels rotation is assumed to be in a sun-synchronous configuration mode. The panels contain surface-bonded piezoelectric patches that can be used either as sensors for the elastic displacements or as actuators to counteract the vibration motion. It is assumed that in the normal mode operation the satellite platform points towards the Earth while the solar arrays rotate so as to follow the Sun. The vehicle moves in a low Earth polar orbit. The technique used to obtain the mathematical model combines the Lagrangian formulation with the Finite Elements Method used to describe the dynamics of the solar panel. The gravity-gradient torque as well as the torque due to the interaction of the Earth magnetic field and the satellite internal residual magnetic moment is included as environmental perturbations. The actuators are three reaction wheels for attitude control and piezoelectric actuators to control the flexible motion of the solar arrays. Computer simulations are performed using the MATLAB® software package. The following on-orbit satellite operating configurations are object of analysis: i) Satellite pointing towards the Earth (Earth acquisition maneuver) by considering the initial conditions in the elastic displacement equal to zero, aiming the assessment of the flexible modes excitation by the referred maneuver; ii) the satellite pointing towards the Earth with the assumption of an initial condition different from zero for the flexible motion such that the attitude alterations are checked against the elastic motion disturbance; and iii) attitude acquisition accomplished by taking into account initial conditions different from zero for both attitude and elastic vibrations. Additionally, the control efforts for the three cases are compared. Results indicate that the attitude control is able to excite the solar panels' vibration modes and vice-versa. The piezoelectric vibration control shows significant performance improvement when compared to contributions of the attitude control to the vibration damping.

© 2017, Brazilian Society for Automatics--SBA.This paper treats the problem of position formation flight control of a group of three multirotor aerial vehicles under obstacle and collision avoidance constraints. In order to solve the problem, a distributed architecture with model predictive controllers for each vehicle includes a set of convex constraints on the vehicles’s position to prevent collisions with other vehicles and obstacles. The resulting distributed scheme controls the formation based on a virtual structure approach where the computers of the architecture exchange position data through diagrams in Simulink. The performance of the method is assessed through simulations considering that the vehicles are subject to disturbance forces and the results show the effectiveness and the ability of the control architecture to handle the obstacle and collision avoidance constraints.

© 2017 IAAIn space the manipulators working space is characterized by the microgravity environment. In this environment the spacecraft floats and its rotational/translational motion may be excited by any internal and external disturbances. The complete system, i.e., the spacecraft and the associated robotic manipulator, floats and is sensitive to any reaction force and torque related to the manipulator's operation. In this sense the effort done by the robot may result in torque about the system center of mass and also in forces changing its translational motion. This paper analyzes the impact of the robot manipulator dynamics on the attitude motion and the associated control effort to keep the attitude stable during the manipulator's operation. The dynamics analysis is performed in the close proximity phase of rendezvous docking/berthing operation. In such scenario the linear system equations for the translation and attitude relative motions are appropriate. The computer simulations are implemented for the relative translational and rotational motion. The equations of motion have been simulated through computer by using the MatLab software. The LQR and the PID control laws are used for linear and nonlinear control, respectively, aiming to keep the attitude stable while the robot is in and out of service. The gravity-gradient and the residual magnetic torque are considered as external disturbances. The control efforts are analyzed for the manipulator in and out of service. The control laws allow the system stabilization and good performance when the manipulator is in service.

© 2017 Elsevier LtdThe objective of this paper is to describe a new method for modeling aeroelastic system in time domain based on a modification of the Laguerre Polynomials to represent complex quantities. These polynomials are used to approximate the unsteady aerodynamics forces which are defined in the frequency domain using the Doublet Lattice Method (DLM). In this approach, the size of the matrices representing the aeroelastic system remains the same as the matrices representing the structural dynamics behavior. It is an important point since classical state space aeroelastic models include lag states increasing the size of the matrices used to represent the system. The applicability of the method is demonstrated by numerical simulation performed on the benchmark wing structure. The approach offers promise mainly for complex systems such as real aircraft.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This paper presents the algorithms for autonomous modes for a tether airship. Switching between flight modes, the airship can perform different flight tasks as such as autonomous hovering, hovering aligned to the home point, take-of and waypoint navigation. The switch-ing scheme between flight modes and the designed control loop are also included in this paper. The control loop is divided into two loops, the inner and the outer control loop. The stability augmentation system has been included as inner control loop to stabilize longitudinal and lateral dynamics and for cancellation of external disturbance. The outer control loop generates control inputs depending of flight mode and given references. For each flight task it is performed outdoor flight experiment with a small size tethered airship, and on the results, the tethered airship performance is shown.

© 2017, American Institute of Aeronautics and Astronautics. All rights reserved.A preliminary investigation of applying system identification methodology for a tethered airship vehicle is presented. A series of flight test campaigns was carried out with the tethered airship providing flight data suitable for parameter estimation. A nonlinear model is selected and a set of state and observation equations is described; biases, and initial condition parameters are estimated by data compatibility check and the flight path reconstruction is presented. The Output Error Method is selected to estimate the aerodynamic parameters and the initial conditions taking into account the influence of tether disturbances. This paper also presents the experiment setup, flight conditions, and brings preliminary system identification results regarding the tethered airship dynamics. Finally, it brings a discussion regarding main considerations on approaching system identification methods for a tethered airship.

Copyright © (2017) by International Astronautical Federation. All rights reserved.This paper presents the mathematical modeling of a 5 degree-of-freedom robot manipulator-like spacecraft and the computer simulations of the attitude motion in a low earth orbit. The main goal of the study is the study the effects of the space vehicle center of mass changes in the rotational motion. The center of mass moves due to changes in the mass configuration of the robot arm during the robot manipulator orbital operations. In such analysis it is not reasonable to assume the inertia matrix diagonal since the motion of the robot links causes the appearing of products inertia. In the same way it is recommended to consider the non linear equations of motion. Another feature of the work is that the rotational motion control aims to suppress the reactions forces and torques in the manipulator joints. The reaction forces affect the translational and the attitude motion as well. The results show that the control of the translational and attitude motions keeps the dynamics as planned, showing and suppressing the center of mass motion when the manipulator is operating.

© 2016 The Authors.This work presents experimental results of two damage detection techniques based on modal properties, with the application on a full-size composite helicopter main rotor blade. The damage detection methods used in this study are the coordinate modal assurance criterion (COMAC) and the modal strain energy method, which are respectively based on the comparison of vibration modes and on the comparison of the modal strain energy of a beam. Modal parameters were obtained with experimental modal analysis and damage was introduced artificially on the blade by attaching a small mass to it, changing its global properties in this way. Finally, experimental results for the damage detection technique are shown for both methodologies, and remarks concerning sensitivity and robustness of the methods are discussed.

© 2016, Journal of Aerospace Technology and Management. All rights reserved.The Operational Modal Analysis technique is a methodology very often applied for the identification of dynamic systems when the input signal is unknown. The applied methodology is based on a technique to estimate the Frequency Response Functions and extract the modal parameters using only the structural dynamic response data, without assuming the knowledge of the excitation forces. Such approach is an adequate way for measuring the aircraft aeroelastic response due to random input, like atmospheric turbulence. The in-flight structural response has been measured by accelerometers distributed along the aircraft wings, fuselage and empennages. The Enhanced Frequency Domain Decomposition technique was chosen to identify the airframe dynamic parameters. This technique is based on the hypothesis that the system is randomly excited with a broadband spectrum with almost constant power spectral density. The system identification procedure is based on the Single Value Decomposition of the power spectral densities of system output signals, estimated by the usual Fast Fourier Transform method. This procedure has been applied to different flight conditions to evaluate the modal parameters and the aeroelastic stability trends of the airframe under investigation. The experimental results obtained by this methodology were compared with the predicted results supplied by aeroelastic numerical models in order to check the consistency of the proposed output-only methodology. The objective of this paper is to compare in-flight measured aeroelastic damping against the corresponding parameters computed from numerical aeroelastic models. Different aerodynamic modeling approaches should be investigated such as the use of source panel body models, cruciform and flat plate projection. As a result of this investigation it is expected the choice of the better aeroelastic modeling and Operational Modal Analysis techniques to be included in a standard aeroelastic certification process.

© 2015, The Brazilian Society of Mechanical Sciences and Engineering.The acoustic design of cavities is an important task in a variety of engineering applications, from automotive or aerospace industries to equipment coating designs. In this work, the acoustic impedance functions (a frequency domain model) were calculated using analytical, numerical, and experimental methods. Those different approaches were presented in a unified manner in order to allow comparisons among them. The relationship of the impedance function and a classical frequency response function (FRF) was also established. A circular duct of rigid walls was assumed with different boundary conditions as closed end, as well as opened and absorbed extremities. Three duct configurations were implemented in order to compare analytical, numerical, and experimental results. Finally, it could be possible to evaluate some aspects that are characteristic of a large range of acoustic systems applications as the existence of complex modes and frequency-dependent behavior of absorption material. This study aims the usage of the impedance functions to analyze the acoustic behavior of cavities, as well as to compose the background in order to develop, in the future, an acoustic modeling process using impedance functions.

Complex modes are commonly observed in systems where damping acts locally, i.e., in cases the damping distribution is not uniform. The present work aims to compare the natural modes resultant from different boundary conditions of a cavity, considering a circular duct of rigid walls as the cavity under test. In one end, a volume velocity source was installed. At the other end, were considered four cases: closed end, open end, and two absorbing ends with different thickness. An Acoustic Experimental Modal Analysis was performed, in order to evaluate the mode shapes, using as basis a set of Acoustic Impedance Functions. Quantitative (phase scatter) and qualitative (variation of the node location) analysis were performed to estimate the "complexity" of the mode shapes. The comparison among the resultant complex mode shapes for each boundary condition showed the relation between the level of resistance and its effect in the complex modes.

The most common and established way of performing experimental modal analysis is to use acceleration based transducers that lead to the calculation of the displacement mode shapes. However, the use of strain measurements for experimental modal analysis has recently gained a lot of popularity. Not only there are applications where the use of strain measurements makes for a more attractive and interesting option, such as structural health monitoring methods, but there are also applications where sensor size and placement might be critical and therefore strain sensors are the most eligible candidate. This work has as the main focus of research the use of strain sensors for experimental modal analysis. In this sense, experimental methodologies and improvements on the current ways of carrying out strain modal analysis are presented, paying particular attention to the relationship between strain and displacement modes. This study of the strain displacement relationship led to the development of a scaling methodology for strain modes and is used to demonstrate the presence of reciprocity under certain conditions. To validate the proposed methodologies, beam and plate structures will be analyzed and experimental results will be shown.

Copyright © 2016 by ASME.Increasingly, the product development industry coexists with highly complex and integrated systems, parallel to the search for engineering tools that are not difficult to use, reduce the complexity and effort to successfully model the systems and, at the same time, provide reliable results. In this context, to demonstrate the differences between the signal-port and the multi-port modeling approaches, and to analyze the modeling effort demanded by each, this work aims at building, in MATLAB/Simulink® and LMS Imagine. Lab AMESim, a model of flight control's electro-hydrostatic actuator (EHA). Further, this same EHA system is built in a model that integrates both environments. Simulations then provided data for performance comparisons and conclusions about the possible gains of this integration, for industrial research & development purpose. Lastly, the physical nonlinear EHA AMESim model is linearized, rendering a transfer function model, available to be used in linear control studies in MATLAB-Simulink.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This paper presents the dynamic analysis of an unmanned tethered airship in hovering flight using a stability augmentation system. A comparison between tethered aerostats and airships is made, and the benefits and issues related to a tethered airship are highlighted. Since airships are not designed to be stable as tethered aerostats, a stability augmentation system is implemented in the flight control system. The dynamic equations of motion and the controller for the tethered airship are described. A linearized model is obtained by finite difference approach and the gains are obtained using LQR technique. The details of the flight control system and experimental set up are provided. Outdoor flight testing of the tethered airship was conducted and the stability of natural dynamics and active control response was analyzed in the time domain using flight test data. The results obtained validated the closed-loop system for the range of wind condition tested once the stability augmentation is achieved.

© The Society for Experimental Mechanics, Inc. 2016.Experimental modal analysis is commonly associated with the use of simulation models for validation, correlation and model updating. However, this interaction between simulation and test is constantly evolving, not in the least because it can be applied to model-based design engineering in the broad sense. Over time, new simulation methods have emerged and consequently, new approaches combining experimental and numerical methodologies are needed and possible. Model Based System Testing (MBST) is an innovative paradigm that allows to structure this process and, in particular, to investigate how the well-established modal testing and analysis procedures and ways of working can be adopted to the multiphysical nature of mechatronic systems. As a result, many possibilities arise: test data can be used to validate multiphysical models, models help gaining insights into test conditions, hybrid approaches allow combining testing and simulation on hardware-in-the-loop and system-in-the-loop test benches, where physical systems can be combined with simulation models to apply loads and more realistic test conditions, as well as the use of data coming from feedback control system information for testing purposes. In this paper, the context and concepts of MBST will be introduced, and application examples will be shown, highlighting the advantages of such a methodology.

© The Society for Experimental Mechanics, Inc. 2016.Strain modal analysis has been used for a long time as an alternative way of identifying the vibration modes of a structure, whenever the use of accelerometers is not suited (such as in aerospace applications) or when information about the dynamic strain levels in the structure is required (e.g. structural health monitoring or durability applications). However, some practical aspects of strain-based modal analysis are not always studied in such applications. For instance, how to visualize and interpret multi-directional modal strain is still an open topic. Similarly, the optimal way to correlate and distinguish these strain modes is also not usually discussed. This work will focus on clarifying some concepts related to multi-directional strain, and will show some examples on how modal strain in multiple directions can be better interpreted. For this purpose, the use of strain gauges (and strain rosettes) for modal analysis in two-dimensional structures will be introduced and some insights will be given on how to properly measure and interpret normal and shear strain modes and how to carry out modal correlation. Finally. these concepts will be applied to numerical and experimental examples.

© 2015 IOP Publishing Ltd.Piezoelectric shunt damping is a well-known technique to damp mechanical vibrations of a structure, using a piezoelectric transducer to convert mechanical vibration energy into electrical energy, which is dissipated in an electrical resistance. Resonant shunts consisting of a resistance and an inductance connected to a piezoelectric transducer are used to damp structural vibrations in narrow frequency bands, but their performance is very sensitive to variations in structural modal frequencies and transducer capacitance. In order to overcome this drawback, a piezoelectric shunt damping technique with improved performance and robustness is presented in this paper. The design of the adaptive circuit considers the variation of the host structure's natural frequency as a project parameter. This paper describes an adaptive resonant piezoelectric vibration absorber enhanced by a synthetic negative capacitance applied to a shell structure. The resonant shunt circuit autonomously adapts its inductance value by comparing the phase difference of the vibration velocity and the current flowing through the shunt circuit. Moreover, a synthetic negative capacitance is added to the shunt circuit to enhance the vibration attenuation provided by the piezoelectric absorber. The circuitry is implemented using analog components. Validation of the proposed method is done by bonding the piezoelectric absorber on a free-formed metallic shell.

© 2015, Springer Science+Business Media Dordrecht.The common practice in industry is to perform flutter analyses considering the generalized stiffness and mass matrices obtained from finite element method (FEM) and aerodynamic generalized force matrices obtained from a panel method, as the doublet lattice method. These analyses are often re-performed if significant differences are found in structural frequencies and damping ratios determined from ground vibration tests compared to FEM. This unavoidable rework can result in a lengthy and costly process of analysis during the aircraft development. In this context, this paper presents an approach to perform flutter analysis including uncertainties in natural frequencies and damping ratios. The main goal is to assure the nominal system’s stability considering these modal parameters varying in a limited range. The aeroelastic system is written as an affine parameter model and the robust stability is verified solving a Lyapunov function through linear matrix inequalities and convex optimization.

Strain gauges and strain measurements have been widely used in structural health monitoring (SHM) systems as a means of detecting and localizing damage, due to their higher sensitivity to local damage. These damage identification techniques normally use strain related measurements such as the mode curvature, strain frequency response function or strain energy as the main parameter to detect damage. However, damage detection techniques based on acceleration measurements have also been investigated in the past, using modal parameter comparison and other methodologies. In this paper, the use of vibration-based strain measurements for use in SHM systems will be evaluated, with the purpose of characterizing their higher sensitivity in damage detection, when compared to other vibration measurements, such as acceleration-based measurements. Since the choice and use of the most damage sensitive parameter can lead to a more sensitive and robust system, the assessment of the more suitable sensor and processing of information is very important. For this purpose, numerical and experimental examples will be discussed to evaluate the higher performance of the strain gauges.

© 2015 IEEE.Among the main sub-areas covering the cooperative control problem of Unmanned Aerial Vehicles (UAVs), formation flight has attracted great interest and has been widely investigated. The main purpose of the formation flight control is to establish a desired shape of formation for a group of vehicles by controlling the positions of each vehicle. The present paper deals with the problem of position formation flight control of a group of three multirotor helicopters with collision avoidance. In order to solve the problem, we propose a decentralized scheme based on model predictive controllers (MPC) for formation according to a virtual structure approach. For collision avoidance, a set of convex constraints on the vehicle's positions are included. The proposed method is evaluated on the basis of computational simulations considering that the vehicles are subject to disturbance forces. Simulation results show the effectiveness of the method with primary focus on treatment of anti-collision constraints.

Flight simulators are employed by civil and military pilots, as well by engineers, in order to increase the security in training of crew, and to find out the behavior of the aircraft under different operational conditions. However, it is necessary to calibrate the simulator software to have good adherence to real flight. In this process, parameters of the mathematical model of the flight simulation need to be identified, such that the simulation is as close as possible to the real flight dynamic. With appropriated values of these parameters, the simulator will be ready for training or assessing the aircraft dynamics. This can be described as an inverse problem or parameter identification, formulated as an optimization problem. The simulator is designed to represent the dynamics of the helicopter AS355-F2, for testing two types of maneuverswere employed: a sinusoidal input and 3-2-1-1 pulse input. The aerodynamic derivatives estimation methodology is also known as quad-M scheme, since it involves four different processes: Measurement, Maneuver, Model, and Methods of error minimization. The tested helicopter was equipped with the Aydin Vector Data Acquisition System (AVDAS) PCU-816-I, ATD-800 digital recorder The system measures a total of thirty-five different parameters. The calibration of a dynamic flight simulator is achieved by two meta-heuristics: a Genetic Algorithm and a new approach named Multiple Particle Collision Algorithm (MPCA). Preliminary results show a good performance of the employed optimization methods.

© 2015 AIAA American Institute of Aeronautics and Astronautics. All rights reserved.This paper investigates the performance of an airship controller designed to impart autonomous control ability to an airship fitted with three fixed BLDC motors and propellers. An existing airship designed primarily for indoor flights was chosen as a platform for mounting and testing this control system for autonomous operations. The sub-systems of this airship are described, and details of an autopilot and waypoint navigation system are presented. An open source flight controller for UAVs was adapted for airships, and a Hardware-In-the-Loop simulation was carried out to validate it. The performance of this controller was tested in a few outdoor flight tests. It was seen that autonomous flight was possible, including an autonomous takeoff, but since the airship was underpowered, the ability of the airship to maintain the desired flight path was poor. However, this study established that, in principle, it is possible to provide some autonomous capability to an airship by suitably adapting a flight controller designed for UAVs.

The new aspect of the attitude dynamics and control for robot manipulators while performing on-orbit servicing is that they operate on non fixed bases. Even when the robotic manipulator is mounted on a spacecraft, such spacecraft is not fixed in space. The working space is characterized by the microgravity environment. In this environment the spacecraft fluctuates and its rotational motion may be excited by any internal and external disturbances. The complete system, i.e., the spacecraft and the associated robotic manipulator, fluctuate and is sensible to any reaction force and torque inherent to the EVA (extravehicular activities). In this sense all effort done by the robot may result in torque about its center of mass and even result in forces that can cause translational motion of robotic manipulator center of mass. This paper focuses on such scenario by analyzing the impact of the robot arms dynamics on the attitude motion and the associated control effort to keep the attitude motion stable during the manipulator operation. The focus of the dynamics analysis is the close proximity. In such configuration of chaser-target spacecrafts the linear system equations for the translational relative motion is appropriated for the dynamics analysis. The computer simulations are implemented for the relative translational and rotational (attitude motion). The relative attitude equations representing the spacecraft like a robotic manipulator have been simulated through computer by using the MatLab software package. The PID control technique is used to keep the attitude motion stabilized while the robot arms moves its arms under the effect of gravity-gradient and applies forces to execute some hypothetical tasks. The control effort is analyzed to bring about the necessary actuators to accomplish the spacecraft attitude stabilization while providing on-orbit servicing.

Ground Vibration Testing (GVT) of aircraft is a measurement campaign performed in the development process of an aircraft, with the objective of obtaining experimental data of the aircraft to validate and update the structural dynamic models, which can in turn be used to predict important behavior, such as flutter. These measurements are usually carried out using standard accelerometers, which lead to the identification of the displacement mode shapes. However, the use of strain sensors in vibration and modal related applications has recently gained popularity, due to some advantages, such as sensor size and the fact that strain relates directly to stress. On the other hand, interpreting the strain mode shapes can sometimes be more complex, so the use of both strain and acceleration sensors can lead to a more complete and understandable dataset. In this paper, the main results of a GVT campaign on an F-16 aircraft will be shown, where the full aircraft was instrumented with accelerometers and one of the wings was also fully instrumented with dynamic strain sensors. The main results of the test campaign will be shown, where both strain sensor and accelerometer measurements are processed simultaneously, resulting in the strain and displacement mode shapes, respectively, and some characteristics and advantages of carrying out the tests this way will be presented.

© 2015, IFAC (International Federation of Automatic Control) Hosting by Elsevier Ltd. All rights reserved.The present paper deals with the problem of position control of a flying robot type multirotor helicopter with obstacle avoidance. In order to solve the problem, an architecture with model predictive controller (MPC) minimize the tracking error, where are implemented the inclusion of convex constraints on the position vector, making it possible to avoid obstacles with a flexible trajectory. The proposed method is evaluated on the basis of computational simulations considering that the vehicles is subject to disturbance forces. Simulation results show the effectiveness of the method related to the tracking performance with focus on the treatment of obstacle avoidance constraints.

© The Society for Experimental Mechanics, Inc. 2015.This paper discusses the use optical fiber Bragg grating (FBG) strain sensors for structural dynamics measurements and modal analysis. For some industrial applications, the use of strain sensors (combined or not with accelerometers) can bring benefits such as reduced size and weight. In many of these applications, FBG sensors lead the class of new sensor technologies that make dynamic strain measurements more attractive, with additional qualities such as the reduction of cabling, immunity to electromagnetic interference and higher sensor robustness. On the other hand, the main difficulty in the use of this technology is their integration and synchronization with other types of sensors, since their acquisition usually requires a separate specialized measurement unit. This is an important requirement in modal analysis, where synchronization between input and output measurements is a key issue that can directly affect the quality of the data. In this paper, FBG sensors are used in an experimental modal analysis, where their analogue signal is digitalized on the same way as the electrical sensors, guaranteeing synchronization.

This paper presents the design process and measurement results of a shunted piezoelectric isolator, which can be a good compromise between a solely passive and an active isolator. The system will be used for the reduction of the vibration transmission between two idealized panels of a plane fuselage. During the design process, numerical models of both, the panel structure and the shunted isolator, are used in order to derive reduced state-space matrices. Based on modal superposition, the numerical models describe the dynamic behaviour of the components and are integrated into a simulation environment of the holistic system. The required modal data is derived from experimental and numerical modal analyses of the panel. By means of an analytical description, the geometry of the shunted piezoelectric isolator is automatically optimized to defined goal parameters. Applying an impedance-admittance simulation approach, both the isolator and the shunt circuit are modeled. Using this simulation environment, the configuration and the performance of the shunt can be investigated and adjusted. After hardware realization of the shunted isolator, it is examined in a test setup. The results from test measurements are compared to simulation results of the system. Finally, two shunted isolators are placed between a fuselage panel and an ideal mass. Measurement results show the vibration reduction potential of the semi-passive system in addition to the solely passive isolation effect. This paper indicates the feasibility of shunted piezoelectric elements in addition to the passive isolation system. Furthermore an effective preliminary design strategy for the layout of shunted piezoelectric isolators is presented and compared to measurement results.

Copyright © 2014 by Institute of Fundamental Technological Research, Polish Academy of SciencesAircraft have become increasingly costly and complex. Military and civil pilots and engineers have used flight simulators in order to increase safety of flight through the training of crew. It is necessary to calibrate the simulation for simulators to have good adherence to reality, that is, to identify the parameters that make the simulation as close as possible to the actual dynamics. After determining these parameters, the simulator will be ready to be used in human resources training or assessing the aircraft. Parameter identification characterizes the aerodynamic performance of the aircraft and can be formulated as a problem optimization. The calibration of a dynamic flight simulator is achieved by a new meta-heuristic called multiple particle collision algorithm (MPCA). Preliminary results show a good performance of the employed approach.

© The Society for Experimental Mechanics, Inc. 2014.The most common and established way of performing experimental modal analysis is to use acceleration or velocity based transducers that lead to the calculation of the displacement mode shapes. However, there are applications where the use of strain measurements makes for a more attractive and interesting option. For instance, since strain measurements are more directly related to stress, fatigue and failure, strain-based measurement methods can be a good option for structural health monitoring methods and monitoring systems. Moreover, applications where sensor size and placement might be critical are also good candidates for strain-based methods. Helicopters, wind turbines and gas turbines are a good example where strain gauges are more suited for vibration measurements. Additionally, any sort of system that uses strain gauges for static testing can also use the same sensors for dynamic testing without incurring additional sensor costs, which can be very useful in some situations. Some application cases of dynamic strain measurements and dynamic strain modal analysis are shown in this work, with test subjects such as a composite helicopter blade, a small wind turbine blade and a composite beam. Different types of sensors and excitation methods were also used as well as correlation with a computational model.

This paper shows some recent advances on the use of strain-based measurements and mixed strain/acceleration measurements for experimental modal analysis are shown in this work. The most common and established way of performing experimental modal analysis is to use accelerometers that lead to the identification of the displacement mode shapes. Since strain measurements are directly related to stress, fatigue and failure, strain-based measurement methods can be a good option for structural health monitoring systems or durability related measurements. Moreover, applications where sensor size and placement might be critical are good candidates for strain-based methods. Helicopters, wind and gas turbines are good examples where strain gauges are more suited for vibration measurements, since the sensors dont take up as much space when attached to the blade surfaces. The concepts of strain modal analysis, strain frequency response functions (SFRF), strain fields, as well as the identification of strain mode shapes are introduced in this paper. Similarly, the basic theory for strain-based modal analysis is presented, including the fundamental equations for dynamic strain. The similarities and differences between acceleration-based modal analysis and strain-based modal analysis are pointed out and a comparison is made.

The most established way of performing experimental modal analysis is to use acceleration or velocity based transducers that lead to the calculation of the displacement mode shapes. However, there are applications where the use of strain measurements makes for a more attractive and interesting option. Strain gauges have been commonly used for static load testing of mechanical products in the aeronautic, automotive and mechanical industry. Moreover, fatigue testing, durability analysis and lifetime prediction has also been a common application where strain gauges are used. This sort of testing is a common part of the product development process, and additional information on product durability and dynamic performance can be assessed by obtaining the modal parameters of the system, while still using the same instrumentation. Moreover, since strain measurements are more directly related to stress, fatigue and failure, strain-based measurement methods can be a good option for structural health monitoring methods and monitoring systems. Applications where sensor size and placement might be critical are also good candidates for strain-based methods. Helicopters, wind turbines and gas turbines are a good example where strain gauges are more suited for vibration measurements. Some application cases of dynamic strain measurements and dynamic strain modal analysis are shown in this work, with test subjects such as a composite helicopter blade.

Connecting an electrical impedance to a piezoelectric transducer bonded onto a mechanical structure is a popular technique named piezoelectric shunt damping. Resonant shunts consisting of a resistance and an inductance connected to a piezoelectric transducer are used to damp structural vibrations in narrow frequency bands, but their performance is very sensitive to variations in structural modal frequencies and transducer capacitance. This paper describes the design of an adaptation circuit for a resonant piezoelectric shunt enhanced by a synthetic negative capacitance. The resonant shunt adapts its value of inductance autonomously by comparing the phase difference of the vibration velocity and the current flowing through the shunt circuit. A synthetic negative capacitance is added to the shunt circuit in order to enhance the damping performance. Validation of the proposed method is done by measurements with a piezoelectric transducer bonded onto a shell and the circuitry is implemented using analog components.

In this paper, the vibrational aspects of an active Gurney flap system with application on a rotorcraft blade are investigated through wind tunnel tests. The test set-up consisted of a representation of a carbon fiber/composite material helicopter main rotor blade section with full size chord and 0.9 m span, and the active Gurney flap mechanism was located near the trailing edge of the blade, while the actuator was placed on the trailing edge part of the blade. The tests were carried out in multiple attack angles, with different deployment schedules for the Gurney flap. In total, four miniature size accelerometers were used to measure the vibration near the trailing edge. Important aspects regarding the active Gurney flap system were analyzed by measuring the vibration and acceleration.

This paper presents a new methodology to analyze aeroelastic stability in a continuous range of flight envelope with varying parameter of velocity and altitude. The focus of the paper is to demonstrate that linear matrix inequalities can be used to evaluate the aeroelastic stability in a region of flight envelope instead of a single point, like classical methods. The proposed methodology can also be used to study if a system remains stable during an arbitrary motion from one point to another in the flight envelope, i.e., when the problem becomes time-variant. The main idea is to represent the system as a polytopic differential inclusion system using rational function approximation to write the model in time domain. The theory is outlined and simulations are carried out on the benchmark AGARD 445.6 wing to demonstrate the method. The classical pk-method is used for comparing results and validating the approach. It is shown that this method is efficient to identify stability regions in the flight envelope. © 2014 Elsevier Ltd.

This work presents a strategy to control nonlinear responses of aeroelastic systems with control surface freeplay. The proposed methodology is developed for the three degrees of freedom typical section airfoil considering aerodynamic forces from Theodorsen's theory. The mathematical model is written in the state space representation using rational function approximation to write the aerodynamic forces in time domain. The control system is designed using the fuzzy Takagi-Sugeno modeling to compute a feedback control gain. It useds Lyapunov's stability function and linear matrix inequalities (LMIs) to solve a convex optimization problem. Time simulations with different initial conditions are performed using a modified Runge-Kutta algorithm to compare the system with and without control forces. It is shown that this approach can compute linear control gain able to stabilize aeroelastic systems with discontinuous nonlinearities. © 2014 Douglas Domingues Bueno et al.

This paper presents a multiphysics modeling of a switched reluctance motor (SRM) to simulate the acoustic radiation of the electrical machine. The proposed method uses a 2-D finite-element model of the motor to simulate its magnetic properties and a multiphysics mechatronic model of the motor and controls to simulate operating conditions. Magnetic forces on the stator are calculated using finite-element analysis and are used as the excitation on a forced response analysis that contains a finite-element model of the motor stator structure. Finally, sound power levels are calculated using the boundary element method. Simulation results of the model are shown and compared with experimental measurements for a four-phase 8/6 SRM. © 1982-2012 IEEE.

This paper presents a methodology for system prognosis based on indicative parameter time series of the equipment condition. The time series is divided in different candidate scenarios according to modifications on exogenous variables that represent external environmental conditions. Each valid scenario is associated with a specific progression model built based on ARIMA time series analysis approach. The forecast model is determined by merging the current scenario progression model with the progression model associated with most similar past scenario. The feasibility and effectiveness of the approach proposed is demonstrated through the prediction of the deg radation characteristics provided by DC machine benchmark fault simulator. © 2013 IEEE.

This work presents experimental results for structural health monitoring method based on modal properties. The test subject is a composite helicopter main rotor blade, which is a highly flexible, slender beam that can display unusual dynamic behavior with orthotropic properties. A damage method detection based on the coordinate modal assurance criterion (COMAC) was implemented and evaluated on a real-size helicopter main rotor blade. This method uses the global modal properties of the system to identify and locate damaged based on a sensor network. Additionally, simpler methods of tracking changes and detecting damage are evaluated, such as natural frequency tracking and modal assurance criterion (MAC). A test setup was built to carry out an experimental modal analysis on the main rotor blade. For that purpose, a total of 55 uniaxial accelerometers were used for the sensor network along the blade in a way to measure the most significant vibration modes and an electrodynamic shaker was used to excite the system. Vibration modes and natural frequencies were identified by means of a least squares fit (PolyMAX), with damage being simulated on the blade by attaching a small mass to it, changing its global properties this way. Experimental results for the damage detection techniques are shown and a comparison between the methods is also made. © 2012 AIAA.

This work presents experimental results for structural health monitoring method based on modal properties. The test subject is a composite helicopter main rotor blade, which is a highly flexible, slender beam that can display unusual dynamic behavior with orthotropic properties. A damage method detection based on the coordinate modal assurance criterion (COMAC) was implemented and evaluated on a real-size helicopter main rotor blade. This method uses the global modal properties of the system to identify and locate damaged based on a sensor network. Additionally, simpler methods of tracking changes and detecting damage are evaluated, such as natural frequency tracking and modal assurance criterion (MAC). A test setup was built to carry out an experimental modal analysis on the main rotor blade. For that purpose, a total of 55 uniaxial accelerometers were used for the sensor network along the blade in a way to measure the most significant vibration modes and an electrodynamic shaker was used to excite the system. Vibration modes and natural frequencies were identified by means of a least squares fit (PolyMAX), with damage being simulated on the blade by attaching a small mass to it, changing its global properties this way. Experimental results for the damage detection techniques are shown and a comparison between the methods is also made. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Low altitude aerostats are usually uncontrolled and dependent on passive stability solutions-sometimes they do not have any. However, their low speed, the high turbulence and the gusts present in the lower boundary layer, as well as their large surface area bring the stability solution's theme into a foreground. This paper aims to survey common approaches to the stability problem when the aerostat is subjected to turbulent winds and gusts and compare them with new active control techniques, especially for low altitude aerostats. Addressed topics are: the dynamic modeling of tether and aerostat; comparison between different positions of the tether confluence point and analyses of elevator control demands. Conclusions are given to make the use of aerostats more reliable for operations with diverse electronic equipment which have tight requirements for oscillation and altitude changes.

Most of the established procedures for analysis of aeroelastic flutter in the development of aircraft are based on frequency domain methods. Proposing new methodologies in this field is always a challenge, because the new methods need to be validated by many experimental procedures. With the interest for new flight control systems and nonlinear behavior of aeroelastic structures, other strategies may be necessary to complete the analysis of such systems. If the aeroelastic model can be written in time domain, using state-space formulation, for instance, then many of the tools used in stability analysis of dynamic systems may be used to help providing an insight into the aeroelastic phenomenon. In this respect, this paper presents a discussion on the use of Gramian matrices to determine conditions of aeroelastic flutter. The main goal of this work is to introduce how observability gramian matrix can be used to identify the system instability. To explain the approach, the theory is outlined and simulations are carried out on two benchmark problems. Results are compared with classical methods to validate the approach and a reduction of computational time is obtained for the second example. © 2013 Douglas Domingues Bueno et al.

This work presents experimental results for structural health monitoring methods based on modal properties. Two test subjects are considered for the study - a simple composite beam in T-shape and a more complex and realistic system, a composite helicopter main rotor blade (MRB). Different sorts of sensors are used for the purpose of identifying damage, such as piezo strain sensors and accelerometers for the composite T-beam and just accelerometers for the helicopter MRB. Damage is simulated on the test objects by attaching a small mass to different locations. The damage detection methods studied are the coordinate MAC (COMAC), enhanced COMAC (eCOMAC) and scaled COMAC (sCOMAC). The modal properties from the composite T-beam and from the helicopter MRB are obtained using a least squares fit (PolyMAX) and the various techniques mentioned above are implemented and compared.

The use of composites in the aircraft industry has generated a great need for structural health monitoring (SHM) and damage detection systems, to allow for safer use of complex materials. Such is the case with helicopter blades - these components nowadays are mostly composed of carbon fiber or glass fiber reinforced plastics laminates, epoxy and honeycomb filled core structures. The use of composite materials on the main rotor blade (MRB) also allows for more complex and efficient shapes to be designed, but at the same time, their use requires an additional effort when it comes to structural monitoring, since damage can occur and go unnoticed. This work presents experimental results for structural health monitoring method based on strain energy. The test subject is a full-scale composite helicopter main rotor blade, which is a highly flexible, slender beam that can display unusual dynamic behavior with orthotropic behavior. This damage detection method is based on the modal strain properties, and a damage detection index is used to identify and quantify damage. A test setup was built to carry out an experimental modal analysis on the main rotor blade. For that purpose, a total of 55 uniaxial accelerometers were used on the helicopter blade to measure the displacement modes of the structure. To compute the strain modes from the displacement modes, central differences approximation is used. Damage is introduced on the blade by attaching a small mass to two different locations. Experimental results show the possibility of locating damage in this case. © (2013) Trans Tech Publications.

This paper presents a model to simulate the dynamic behavior of cabin temperature in an aircraft in the airline market. The model includes the most important components in Environmental Control System (ECS) mainly the air conditioning packs and packs bypass valves. The model predicts the cabin temperature in the aircraft which may fly in several operating conditions. Some case studies are presented and the results are compared to experimental data collected in a similar aircraft. It was found a good agreement between results predicted by the model and the experimental data. The simulation model may be used to evaluate the cabin temperature control and to improve the components design and the performance of ECS under transient conditions. Based on this analysis the ECS controller may be designed to improve the response time and the cabin temperature control stability. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The objective of this study is to describe the design and the implementation of an experimental set-up used to study the dynamics, the experimental identification, and the active vibration control of a flexible structure mounted manipulator system. The system consists of a three-degree-of-freedom cylindrical manipulator system with a flexible link on its tip. A two-degree-of-freedom polar rigid manipulator is mounted on the flexible macromanipulator. The dynamic modelling and experimental modal analysis identification in the frequency domain are being applied to design active digital control strategies for the micro-manipulator system to damp the mechanical vibrations of the flexible structure on the tip of the macro-manipulator system. © Authors 2011.

This paper presents the application of a method that uses acoustic Frequency Response Functions (FRF) to predict modifications in an original acoustic system: in this case, a cavity of a rocket engine combustion chamber. As a cavity modification, the insertion of Helmholtz Resonators (HR) was applied. The use of HR in such chambers attenuates the effect of combustion instability, which can seriously damage the engine. The Acoustic Modification Prediction (AMP) method using FRF is based on the Structural Modification Using Response Functions (SMURF), which is a well-known structural reanalysis technique. The AMP uses the FRF matrix from the original cavity and the analytical model of the HR to predict the behavior of the new cavity, avoiding the use of large models. In order to validate the prediction results, experimental data were used. The first results presented differences in magnitudes, but it was possible to identify the modification behavior, as well as parameters to be enhanced to have a fully validated method. © (2012) by the Katholieke Universiteit Leuven Department of Mechanical Engineering All rights reserved.

The objective of this paper is the identification of aerodynamic parameters by the use of aeroelastic response using orthogonal functions. The unsteady aerodynamic forces acting on an aeroelastic system can calculated in the subsonic regime by use of the doublet lattice method, for example, providing a model in frequency domain that needs to be coupled structural dynamics. To obtain a time domain representation of the aeroelastic system, rational functions approximation can be used to represent aerodynamic forces. The most common methods found in literature to approximate these unsteady generalized forces from the frequency to time domain are the least square (LS), matrix Padé, and minimum state. In this context, this work proposes the use of orthogonal functions to represent the aeroelastic output from a state space representation of the system. These functions are easily integrated by using a so-called operational matrix of integration. Consequently, it is possible to transform the aeroelastic equations of motion into algebraic equations. After mathematical manipulation the unknown aerodynamic parameters are determined. Numerical simulations, involving a typical section (three degree-of-freedom) aeroelastic system are used to help illustrate the method that could be applied in an identification experiment. © (2012) by the Katholieke Universiteit Leuven Department of Mechanical Engineering All rights reserved.

Recently developed materials and technology make it possible to operate a network of balloons to perform as effectively as a network of satellites for certain communication objectives. Balloons are the only unmanned air vehicle with a specific operation regulation that allows for its use even in the proximity of aircraft routes. We are concerned with the use of a network of balloons to operate a communication platform in high altitude (60 to 100 thousand feet). Such a solution could perform certain tasks as well as a network of satellites, but with a fraction of the investment. In this paper we discuss the use of such network for communication support in certain areas where many off-shore oil fields are located in Brazil. We comment on the most relevant technical and economic aspects of the feasibility of such network. In particular, we discuss: air-space security; winds forecast and choice of best altitude for each balloon to be launched; recovery logistics and optimization of trade-offs between cost and autonomy. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The main purpose of this paper is to study the attenuation of the jump phenomena associated with the Sommerfeld Effect introduced by the nonlinearities of a magnetic rheological damper (MRD) in a non-ideal vibrational system, excited by a DC motor modeled as limited power source. Numerical simulations of the nonlinear vibrations of the system are carried out for different values of the MRD control parameter in order to show the amplitude reduction of the vibrations close to the system resonance introduced by the nonlinear damping effect of the MR system. © The Author(s) 2010.

It was purposed in this study the use of thermal comfort index as feedback parameter for multivariable control of aeronautical air conditioned system. Simulating the developed thermal model, the efficiency gain by using this control law was observed by comparing it with standard control models. Copyright © 2011 SAE International.

The acoustic modes of closed cavities play an important role in the characterization of combustion chamber behavior. On automotive internal combustion engines, an undesired spontaneous ignition phenomenon, called knock, can occur and the detection of this phenomenon can be based on the acoustic modes of the combustion chamber. This is possible, once the high pressure levels inside the combustion chambers excite internal standing waves, causing the lubricant oil dilution and, hence, damaging the liners, pistons and rings. On the other side, in the space industry, combustion instabilities have become a serious problem in gas turbines and dynamics of chambers must be characterized, since the vortical motions can couple directly to an acoustic field, or may produce local acoustic sources if they impinge on a surface. This paper describes a methodology to determine the acoustical longitudinal natural frequencies and modes of a generic cylindrical cavity that is surrounded by a metallic surface. The Finite Element (FE) model was validated through Experimental Acoustic Modal Analysis (EAMA). The Frequency Response Functions (FRFs) were accessed using a volumetric acoustic source and a microphone. The results showed that the numerical natural frequency of each mode agreed with those measured in EAMA. The visual presentation of the experimental extracted modes seems to be poor, compared to the FE results, due to the difficulty of showing the mode shapes using a restricted number of measurement points. However, a reasonable mode characterization can be obtained with EAMA. In addition, a coupled Frequency Response Analysis (FRA) of the referred structural-acoustic system was calculated, in order to evaluate the vibro-acoustic coupling, considering a FE/FE model.

In this paper, the dynamical response of a coupled oscillator is investigated, taking in consideration the nonlinear behavior of a SMA spring coupling the two oscillators. Due to the nonlinear coupling terms, the system exhibits both regular and chaotic motions. The Poincaré sections for different sets of coupling parameters are verified. © 2011 World Scientific Publishing Company.

Copyright © 2010 SAE International.The free-fall operation comprises a redundant, dissimilar and independent mechanically operated method of extending airplane landing gear due to a main hydraulic system failure or an electrical system malfunction. However, the emergency extension operation system design is not unique and spring-assisted, auxiliary hydraulics-assisted or even pneumatics-assisted landing gear free-fall design can be found in different airplanes. This paper aims at describing the model simulation and the optimization of certain parameters related to the associated hydraulic system, for emergency operation condition, in a non-assisted system configuration comprising simple extension by gravity. Since the free-fall modeling involves different subjects like landing gear extension dynamics, hydraulic actuator kinematics, fluid mechanics and even aerodynamic drag, which illustrates the complexity behind its simulation and optimization, a deep literature review was accomplished in order to support all the formulation necessary to make the modeling feasible. For this purpose, a parametric model was created in MATLAB Simulink, which, by means of an iterative process, allowed the determination of specific parameters values that optimized the damping for that operation. Parameters like restrictor orifices and hydraulic actuator piston areas were evaluated for a chosen landing gear configuration and system performance optimized through the assistance of MATLAB optimization tools. Finally, the purpose of the optimum damping comprised the attenuation of the impact effects suffered by aircraft structure when landing gear falls by gravity in an emergency operation, as well as the assurance of sufficient energy for landing gear locking at the end of its downward movement.

SMART material systems offer great possibilities in terms of providing novel and economical solutions to engineering problems. The technological advantages of these materials over traditional ones are due to their unique microstructure and molecular properties. Smart materials such as shape memory alloys (SMA), has been used in such diverse areas of engineering science, nowadays. In this paper, we present a numerical investigation of the dynamics interaction of a nonideal structure (NIS). We analyze the phenomenon of the passage through resonance region in the steady state processes. We remarked that this kind of problem can lead to the so-called Sommerfeld effect: steady state frequencies of the DC motor will usually increase as more power (voltage) is given to it in a step-by-step fashion. When a resonance condition with the structure it is reached, the better part of this energy it is consumed to generate large amplitude vibrations of the foundation without sensible change of the motor frequency as before. The results obtained by using numerical simulations are discussed in details. Copyright © 2009 by ASME.

This paper presents a scheme of multiple neural networks (MNNs) with a new strategy of combination. This combination can obtain an accumulative learning: the knowledge is increased by gradually adding more neural networks to the system. This scheme is applied to flexible link control via feedback-error- learning (FEL) strategy, here called multi-network-feedback-error-learning. Three different neural control approaches are used to control a flexible link, and it is shown that a better inverse dynamic model of the plant is obtained in this case. © 2010 IEEE.

The Active Vibration Control (AVC) technique has been widely used for some kinds of industrial applications, especially in the aeronautical industry, where some manufacturers have applied considerable efforts to offer comfort and weight reduction to aircrafts. This work deals with a numerical implementation of an active multichannel vibration control system that runs over a broadband signal to reduce the vibrations levels of an aeronautical structure (a fuselage section) excited by white noise. The model is based on a real aeronautical fuselage modeled by finite element model, which was verified by experimental modal analysis. The AVC system, based on the Fx-LMS Feedforward algorithm, was implemented for 7 inputs and 6 outputs, using adaptive non-recursive structures (FIR filters) for modeling each structural path and control the displacement at six points of the plant. The results show that the Fx-LMS can be used for broadband excitation signals. Attenuations around 25 dB were obtained in the numerical simulations.

This work presents the results of the modal analysis performed during the ground vibration testing of a testbed originally designed by the Group for Aeronautical Research and Technology in Europe (GARTEUR). The model testing brought challenges in determining modes with very close frequency values, which were detected independently of the excitation signal. A modal validation process was carried out in order to identify these close-spaced modes as well as their dynamic characteristics. The reliability of the experimental modal model was verified by modal assurance criterion calculations between the experimental data and validated by comparison with a finite element model. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

Over the last 40 years, many solid and liquid rocket motors have experienced combustion instabilities. Among other causes, there is the interaction of acoustic modes with the combustion and/or fluid dynamic processes inside the combustion chamber. Studies have been showing that, even if less than 1% of the available energy is diverted to an acoustic mode, combustion instability can be generated. On one hand, this instability can lead to ballistic pressure changes, couple with other propulsion systems such as guidance or thrust vector control, and in the worst case, cause motor structural failure. In this case, measures, applying acoustic techniques, must be taken to correct/minimize these influences on the combustion. The combustion chamber acoustic behavior in operating conditions can be estimated by considering its behavior in room conditions. In this way, acoustic tests can be easily performed, thus identifying the cavity modes. This paper describes the procedures to characterize the acoustic behavior in the inner cavity of four different configurations of a combustion chamber. Simple analytical models are used to calculate the acoustic resonance frequencies and these results are compared with acoustic natural frequencies measured at room conditions. Some comments about the measurement procedures are done, as well as the next steps for the continuity of this research. The analytical and experimental procedures results showed good agreement. However, limitations on high frequency band as well as in the identification of specific kinds of modes indicate that numerical methods able to model the real cavity geometry and an acoustic experimental modal analysis may be necessary for a more complete analysis. Future works shall also consider the presence of passive acoustic devices such as baffles and resonators capable of introducing damping and avoiding or limiting acoustic instabilities.

SMART material systems offer great possibilities in terms of providing novel and economical solutions to engineering problems. The technological advantages of these materials over traditional ones are due to their unique microstructure and molecular properties. Smart materials such as shape memory alloys (SMA), has been used in such diverse areas of engineering science, nowadays. In this paper, we present a numerical investigation of the dynamics interaction of a nonideal structure (NIS). We analyze the phenomenon of the passage through resonance region in the steady state processes. We remarked that this kind of problem can lead to the so-called Sommerfeld effect: steady state frequencies of the DC motor will usually increase as more power (voltage) is given to it in a step-by-step fashion. When a resonance condition with the structure it is reached, the better part of this energy it is consumed to generate large amplitude vibrations of the foundation without sensible change of the motor frequency as before. The results obtained by using numerical simulations are discussed in details. © 2009 by ASME.

This paper investigates an individual-blade-root-control approach to the reduction of helicopter blade sailing and suppression of tunnel strikes for articulated rotors, considering steady flow conditions during engagement shipboard operations. The aeroservoelastic modeling includes a nonlinear structural dynamics related to the droop and flap stops, a linear aerodynamic model based on the blade-element theory, a linear gust model for the ship airwake, and a lift compensator. The blade-sailing model is a forced parametric flapping oscillator with nonlinear stiffness and time-varying coefficients. The aeroelastic control law design yields a flap-state-feedback individual-blade- root controller for the lift/angle-of-attack compensation whose parameters are associated with the damping/stiffness enhancement of the flapping oscillator. The simulation results show that the proposed active aeroelastic controller yields blade-sailing reduction of nearly 30% in upward and downward deflections at severe wind-over-deck conditions by using low blade pitch input limits of the actuators. This blade-sailing reduction can prevent tunnel strikes from occurring. The relaxation of the actuator limits can significantly improve the attenuation of the blade deflections. Copyright © 2009 by the American Helicopter Society International, Inc. All rights reserved.

In this work, two optimization methods are investigated to accomplish the parameter identification of the longitudinal motion of a military aircraft within the framework of the Output Error methodology. One of these methods is based on natural algorithms, in particular, the genetic algorithms and particle swarm optimization, which combined constitute the so called Life Cycle Method. The other is the Levenberg-Marquardt optimization algorithm, which is a gradient based method. Since the methodologies differ in the way they perform the optimization, being the first based on search and so appropriate for global minima search and the second gradient based, thus, good for local minima, both are compared and used in such a way that they complement each other. © 2009 IFAC.

This work deals with the optimization of flight-test maneuvers for aerodynamic parameter estimation considering that the measurements are contaminated with colored residuals. The colored residuals consideration is important to give a direct and realistic assessment of the parameter estimation uncertainty levels before flight testing. The design technique is based on the optimization of the flight-test data information content and the Cramer-Rao lower bound. The discrete autocorrelation matrix of the measurement residuals is also used to compose the optimization criterion, thereby explicitly considering colored residuals. To validate the proposed technique, a flight-test campaign of the CEA-205 CB-9 Curumim was performed and its results discussed. The advantages of the proposed maneuver optimization technique are presented, stressing the ease of implementation of the signals and the strong improvement in the estimation procedures made possible with the application of the optimized maneuver signals. © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

In this work, two optimization algorithms are investigated to accomplish the parameter identification of the longitudinal motion of a real aircraft by using the output error method. The first algorithm is the nature-inspired algorithm named the life cycle model, which is a composed strategy based on other heuristics such as genetic algorithms and particle swarm optimization. The second one is the gradient-based technique named Levenberg-Marquardt algorithm, which is a variant of the Gauss-Newton method. Flight test data, performed with a training jet aircraft (Xavante AT-26), were used to feed the output error method. In this context, both optimization algorithms were tested, in solo performance and in a cascade-type approach. Results are reported, aiming to illustrate the success of using the proposed methodology.

This work describes the application of the output-only modal analysis to the study of the structural dynamic parameters of the Brazilian Air Force F-5E aircraft, in several flight conditions. The applied methodology is based on a technique to estimate the frequency response functions and extract the modal parameters using only the structural dynamic response data, without assuming the knowledge of the excitation forces. The aircraft was multiply excited in flight with impulsive forces from pyrotechnical devices known as bonkers. These devices are attached to different position along the airframe. The in-flight structural response has been acquired by accelerometers distributed along the wings, fuselage and empenages of the aircraft. The Enhanced Frequency Domain Decomposition (EFDD) technique was chosen to identify the dynamic parameters of the airframe. This technique is based on the hypotheses that the system is randomly excited with a broad band spectrum with almost constant power spectral density. The system identification procedure is based on the single value decomposition (SVD) of the power spectral densities of system output signals, estimated by the usual Fast Fourier Transform (FFT) method. This procedure has been applied to different flight conditions to evaluate the modal parameters and the aeroelastic stability trends of this airframe. The experimental results obtained by this methodology were compared with the predicted results supplied by aeroelastic numerical models in order to check the consistency of the proposed output-only methodology. © 2008 SAE International.

This work deals with the application of a nature-inspired optimization technique to solve an inverse problem represented by the identification of an aircraft landing gear model. The model is described in terms of the landing gear geometry, internal volumes and areas, shock absorber travel, tire type, and gas and oil characteristics of the shock absorber. The solution to this inverse problem can be obtained by using classical gradient-based optimization methods. However, this is a difficult task due to the existence of local minima in the design space and the requirement of an initial guess. These aspects have motivated the authors to explore a nature-inspired approach using a method known as LifeCycle Model. In the present formulation two nature-based methods, namely the Genetic Algorithms and the Particle Swarm Optimization were used. An optimization problem is formulated in which the objective function represents the difference between the measured characteristics of the system and its model counterpart. The polytropic coefficient of the gas and the damping parameter of the shock absorber are assumed as being unknown: they are considered as design variables. As an illustration, experimental drop test data, obtained under zero horizontal speed, were used in the non-linear landing gear model updating of a small aircraft.

State and parameter estimation using flight test data is highly affected by process and measurement noises, especially with noises displaying time varying statistical properties. Hence, if an estimation problem is to be solved, an adaptive filtering approach is recommended. It is also desirable to obtain the estimates online, simultaneously with flight execution, aiming at a maneuver validation before concluding the flight. Indeed, it is more expensive to put the aircraft back in the air than to extend a little the flight and repeat a test point. Flight path reconstruction is a technique which produces a consistent flight test data set from noisy measurements as a preprocessing scheme to a parameter identification routine. Air data can also be calibrated simultaneously if the problem is formulated properly. This work proposes a methodology to deal with time varying noise statistical properties using a new approach for an adaptive extended Kalman filter. Besides the main filter, two other Kalman filters are proposed to run in parallel, to estimate the process and measurement noise statistics based on the main filter residuals. The proposed adaptive method is derived from the covariance matching technique, by employing filter residuals to adjust the noise statistical properties. Because the method has a low computational cost and is recursive, it is suitable for online applications. The method is validated in a flight path reconstruction application, with simultaneous air data calibration for angle of attack, angle of sideslip, and static pressure sensors. A 100 samples Monte Carlo simulation and real flight test data analysis are used for performance evaluation. Because the proposed approach adequately estimates the statistical noise properties, improved performance is obtained. Copyright © 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

This work deals with the optimization of flight test maneuvers for aerodynamic parameter estimation considering that the measurements are contaminated with colored residuals. The colored residuals consideration is important to give the direct and realistic assessment of the parameter estimation uncertainty levels prior to flight tests. The optimization technique is based on the concept of flight test data information content and Cramer-Rao lower bound. The discrete autocorrelation matrix of the measurement noise is used in order to compose the optimization criteria considering colored residuals. Some results of a flight test campaign of the CEA-205 CB.9 Curumim aircraft are discussed. The advantages of the proposed maneuvers optimization technique are presented, stressing the easiness of implementation of the signals and the strong improvement in the estimation procedures made possible with the application of the optimized maneuver signals.

Certification requirements, optimization and minimum project costs, design of flight control laws and the implementation of flight simulators are among the principal applications of inverse problem applications in the aeronautical industry. The problem of aircraft identification and parameter estimation demands for accurate mathematical model of the aerodynamics and adequate experimental flight data gathering and processing. The aircraft dynamic modeling is characterized by aerodynamic and control derivatives whose values can be directly determined from flight test data. This work describes the application of the output-error method using the Nelder-Mead (NM) and Levenberg-Marquardt (LM) algorithms to obtain the aerodynamic and control derivatives of a regional jet aircraft. Unlike others identification methods based on equation-error the output-error method gives unbiased parameter estimation in the presence of measurement noise. In this work, experimental results for estimation of the lateral directional aerodynamic derivatives, using flight test data provided by EMBRAER, are presented.

This work deals with the optimization of flight test maneuvers for aerodynamic parameter estimation considering that the measurements are contaminated with colored residuals. The colored residuals consideration is important to give the direct and realistic assessment of the parameter estimation uncertainty levels prior to flight tests. The optimization technique is based on the concept of flight test data information content and Cramer-Rao lower bound. The discrete autocorrelation matrix of the measurement noise is used in order to compose the optimization criteria considering colored residuals. Some results of a flight test campaign of the CEA-205 CB.9 Curumim aircraft are discussed. The advantages and disadvantages of theproposed maneuvers optimization technique are presented, stressing the easiness of implementation of the signals and the strong improvement in the estimation procedures made possible with the application of the optimized maneuver signals.

This work describes the application of the output-error method using the Levenberg-Marquardt optimization algorithm to the Flight Path Reconstruction (FPR) problem, which constitutes an important preliminary step towards the aircraft parameter identification. This method is also applied to obtain the aerodynamic and control derivatives of a regional jet aircraft from flight test data with measurement noise and bias. Experimental results are reported, employing a real jet aircraft, with flight test data acquired by smart probes, inertial sensors (gyrometers and accelerometers) and Global Positioning Systems (GPS) receivers. © 2006 - IOS Press and the authors. All rights reserved.

The terrestrial microwave telemetry links show limitation due their inherent features concerning bandwidth availability, frequency allocation and range. Also it supports only one aircraft per test and the data acquisition capacity can be enhanced. Following the flight tests trends, it proposes a telemetry link based on satellite communications deployed by off the shelf equipments allowing advantages as bandwidth availability, multiple aircrafts telemetry and almost global range into the reliability standards. By simple equations and typical flight tests data it demonstrates the feasibility of the telemetry system proposed for time and costs reduction to optimize flight tests programs.

Fluid-structure interactions are always present in real life dynamic systems, during operations. However, analysts, due to the complexity of building a vibro-acoustic model and also because sometimes the coupling effect has no significance, often apply "one way" analysis, or uncoupled analysis. This analysis procedure is done in two steps, on which the structural part and the fluid part are modeled separately. Nevertheless, sometimes it is important to consider the mutual influence of the vibro-acoustic system, where the acoustic and structural matrices are coupled and the influence of the structural displacement on the fluid domain, as well as the acoustic pressure of the fluid domain on the structural body, are accounted in one coupled matrix. Low frequency coupling techniques, or deterministic coupling techniques, were used to calculate the Brazilian Satellite Launcher Vehicle (VLS) fairing behavior. The well-known structural FEM/fluid FEM technique and structural FEM/fluid BEM technique were applied to model the fairing body and its acoustic cavity. Calculations were done, which yielded the low frequency acoustic cavity as well as the skin responses. This paper describes the applied procedures to build up the vibro-acoustic models of the VLS. The obtained results are described and a comparison between FEM/FEM versus FEM/BEM techniques is presented. In such a comparison, parameters as computational efficiency, allocated memory, processing time, obtained results and modeling are considered.

It is common at the great aerospace companies or research centers to perform flight tests in facilities deployed for aircrafts and systems evaluation. These resources require high planning and investments comprising telemetry system, airborne and terrestrial equipments, calibration labs and specialized staff. It proposes in flight data acquisition means for support tests programs of aircrafts and airborne systems of low complexity in new developments. Also, it can be used as an optimization tool by introducing the telemetry efficiency concept applied to the development of new aircrafts where there are at least two prototypes read to fly.Copyright © 2005 Society of Automotive Engineers, Inc.

This paper describes the dynamical behavior of a nonlinear flexible beam (cubic nonlinearities considered) connected to a dc motor (responsible for the slewing motion) when the angular displacement of the slewing axis and its derivatives are considered to be of a harmonic type and the system is excited near a resonance (present due to the nonlinear contribution). Copyright © 2005 by ABCM.

The objective of this work is to describe the design and the implementation of an experiment to study the dynamics and the active control of a slewing multi-link flexible structure. The experimental apparatus was designed to be representative of a flexible space structure such as a satellite with multiple flexible appendages. In this study we describe the design procedures, the analog and digital instrumentation, the analytical modeling together with model validation studies carried out through experimental modal testing and parametric system identification studies in the frequency domain. Preliminary results of a simple positional control with collocated sensor and actuator is also described. Copyright © 2003 by the International Astronautical Federation. All rights reserved.

The objective of this work is to describe the design and the implementation of an experiment to study the dynamics and the active control of a slewing multi-link flexible structure. The experimental apparatus was designed to be representative of a flexible space structure such as a satellite with multiple flexible appendages. In this study we describe the design procedures, the analog and digital instrumentation, the analytical modeling together with model validation studies carried out through experimental modal testing and parametric system identification studies in the frequency domain. Preliminary results of a simple positional control with collocated sensor and actuator is also described. Copyright © 2003 by the International Astronautical Federation. All rights reserved.

This work deals with nonlinear dynamic modeling of a flexible slewing structure. The interaction between the slewing motion and the structure deflection has been modeled taking in account the nonlinear curvature of the flexible beam. Two different regimes of this nonlinear dynamical system are studied. Firstly, we present the numerical simulations of a nonlinear flexible beam excited near a superharmonic resonance by a prescribed harmonic angular displacement of the hub. The influence of the damping in the frequency response curves is studied and its characteristic jump phenomena investigated. Secondly, a linear approximation is used to interpret the experimental results involving modal identification and real time digital control of the slewing motion. The experimental results are compared with theoretical predictions based on constrained and unconstrained modal expansion of the slewing beam, and neglecting the nonlinear curvature effects. Copyright © 2003 by the International Astronautical Federation. All rights reserved.

During flight missions, space vehicles are subjected to a severe dynamic pressure loading when their rocket-propulsion systems are operated. This loading may be critical for the vehicle components, as well as for the payload such as satellites, which are usually very soft structures. The success of a satellite launching is determined, amongst other measures, by the satellite resistance to the fairing internal acoustic pressure. This paper describes a numerical analysis of the dynamic response of the mechanical structure and the fairing inner acoustic cavity of the Brazilian Vehicle Satellite Launcher (VLS). Finite Element (FE) and Boundary Element (BE) methods are used for the low-frequency analysis with emphasis on the vibro-acoustic coupling effects between the fairing structural vibrations and the inner cavity acoustic pressures. The high-frequency vibro-acoustic behaviour is analyzed using a Statistical Energy Analysis (SEA) model.

Stability and dynamic performance of liquid-propellant rocket engines (LPRE) are two of the fundamental issues in the engine-vehicle integration process. This analysis requires the construction of a detailed model, trying to capture the most realistic phenomena involved, which generally include several sources of uncertainties. In this paper, a methodology for robust modeling and stability analysis is presented. Firstly, the linear models of the LPRE components are obtained by modeling the various physical processes, at a nominal regime of operation. Afterwards, the Laplace transform is applied to derive a block diagram representation of the linear LPRE. The stability study and dynamic analysis are carried out taking in account the uncertainties in parameters of the plant. The robust stability is assured via the Generalized Kharitonov's Theorem; and the robust frequency and step responses are obtained with the use of specialized MATLAB toolboxes. The robust performance of the system in the time domain is obtained in terms of the response to step function input, while taking into account the plant uncertainties, also known as robust step response. A practical application is illustrated by analyzing a simple pressure-fed LPRE system. © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The objective of this work is to describe the design and the implementation of an experiment to study the dynamics, the experimental identification and the active vibration control of a Flexible Structure Mounted Manipulator System (FSMS). The system consists of a three degree of freedom cylindrical manipulator system with a flexible link on its tip. A two-degree of freedom micro-manipulator is mounted on the flexible link of the macro manipulator. The dynamic modeling and experimental modal analysis identification in the frequency domain are being applied to design active digital control strategies for the micro-manipulator system to damp the mechanical vibrations of the flexible structure on the tip of the macro-manipulator system.

This paper discusses the design of an Active Feel System for a Flight Simulator. The objective is to identify relevant aspects of the active feel system design, and to propose guidelines for further analysis required for the selection of the best solution for this kind of system. The characteristics of an active feel system is presented, and the need for such system in a flight simulator is discussed. In addition, some relevant aspects of the implementation of an active feel system in real airplanes are briefly discussed. The design phases discussed in this paper include: definition of the design requirements, proposal of a possible solution, and system preliminary dynamic modeling with Bond-Graphs. Copyright © 2000 Society of Automotive Engineers, Inc.

This paper discusses two approaches for neural control of a flexible link using the feedback-error-learning technique. This technique aims to acquire the inverse dynamics model of the plant and uses a neural network acting as an adaptive controller to improve the performance of a conventional non-adaptive feedback controller. The non-collocated control of a flexible link is characterized as a non-minimum phase system, which is difficult to be controlled by most control techniques. Two different neural approaches are used in this paper to overcome this difficulty. The first approach uses a virtual re-defined output as one of the impacts for the neural network and feedback controllers, while the other employs a delayed reference input signal in the feedback path and a tapped-delay line to process the reference input before presenting it to the neural network. © 2000 IEEE.

Stability and dynamic performance of liquid-propellant rocket engines (LPRE) are two of the fundamental issues in the engine-vehicle integration process. This analysis requires the construction of a detailed model, trying to capture the most realistic phenomena involved, which generally include several sources of uncertainties. In this paper, a methodology for robust modeling and stability analysis is presented. Firstly, the linear models of the LPRE components are obtained by modeling the various physical processes, at a nominal regime of operation. Afterwards, the Laplace transform is applied to derive a block diagram representation of the linear LPRE. The stability study and dynamic analysis are carried out taking in account the uncertainties in parameters of the plant. The robust stability is assured via the Generalized Kharitonov's Theorem; and the robust frequency and step responses are obtained with the use of specialized MATLAB toolboxes. The robust performance of the system in the time domain is obtained in terms of the response to step function input, while taking into account the plant uncertainties, also known as robust step response. A practical application is illustrated by analyzing a simple pressure-fed LPRE system. © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The objective of this work is to describe the design and the implementation of an experiment to study the dynamics, the experimental identification and the active vibration control of a Flexible Structure Mounted Manipulator System (FSMS). The system consists of a three degree of freedom cylindrical manipulator system with a flexible link on its tip. A two-degree of freedom micro-manipulator is mounted on the flexible link of the macro manipulator. The dynamic modeling and experimental modal analysis identification in the frequency domain are being applied to design active digital control strategies for the micro-manipulator system to damp the mechanical vibrations of the flexible structure on the tip of the macro-manipulator system. © 2004 ASCE.

This work describes a lumped parameter mathematical model for the prediction of transients in an aerodynamic circuit of a transonic wind tunnel. Control actions to properly handle those perturbations are also assessed. The tunnel circuit technology is up to date and incorporates a novel feature: high-enthalpy air injection to extend the tunnel's Reynolds number capability. The model solves the equations of continuity, energy and momentum and defines density, internal energy and mass flow as the basic parameters in the aerodynamic study as well as Mach number, stagnation pressure and stagnation temperature, all referred to test section conditions, as the main control variables. The tunnel circuit response to control actions and the stability of the flow are numerically investigated. Initially, for validation purposes, the code was applied to the AWT ('Altitude Wind Tunnel' of NASA-Lewis). In the sequel, the Brazilian transonic wind tunnel was investigated, with all the main control systems modeled, including injection.

Modeling of a robotic manipulator composed of a single flexible joint is presented. The system included a flexible aluminum beam joined by the base to a rotationally inert rigid cube. A potentiometric sensor and tachometer indicated the position and angular velocity of the rigid body. An extensometric bridge-type deformation sensor and a piezoelectric accelerometer were used to measure elastic deformations in the structure, as well as its different degrees of vibrational freedom. A Lagrangian method was used to model the system applying the extended Hamilton principle to obtain equations for the movement of the beam and its environmental conditions. Discretization of the system was carried out using the `Assumed Modes' method, considering the three primary vibrational modes of the beam. The model proved to be representative of the actual dynamical system.

The objective of this work is to describe the design and the implementation of an experiment to study the dynamics and the active control of a slewing multi-link flexible structure. The experimental apparatus was designed to be representative of a flexible space structure such as a satellite with multiple flexible appendages. In this study we describe the design procedures, the analog and digital instrumentation, the analytical modeling together with model validation studies carried out through experimental modal testing and parametric system identification studies in the frequency domain. Preliminary results of a simple positional control where the sensor and the actuator are positioned physically at the same point is also described.

This paper presents the theorical model of an electrohydraulic servovalve, obtained by means of bond graf[ILLEGIBLE] and its experimental identification in order to have the validation of its simulated theoretical model via ARX (auto regressive exogene) parametric model. © Copyright 1993 Society of Automotive Engineers, Inc.