© 2024, International Council of the Aeronautical Sciences. All rights reserved.With the development of electronics and programming in recent years, the possibility of aeronautical projects is being studied by academia and industry, with the aim of improving and adapting them to different projects for new applications and realities. One of these cases is the adaptation of projects such as the ground effect vehicles developed by the Soviet Union during the Cold War. This is an aircraft capable of flying close to the surface of the water and whose advantage is the energy saving of the propulsion of up to 40%. There are several companies developing projects of this type around the world, adapting them to the capacity and operating conditions of the different realities. In Brazil, the startup Aeroriver is developing a ground effect vehicle, the Volitan. This project aims to improve the transportation of people and cargo on the rivers of the Amazon. For the project to be successful, it is necessary to know up to what altitude this aircraft can fly to demonstrate energy savings, safety and maneuverability. A sub-scale prototype has been developed for initial testing and is currently being tested to determine the range and flight efficiency improvement of the Volitan in ground effect. Propulsion is provided by electric motors and power is supplied by a battery bank, allowing 15 minutes of flight autonomy. In this paper, the development of the electronics and instrumentation of a prototype is presented. In order to measure the efficiency of Volitan in flight, it will be equipped with load cells to measure the thrust force, RPM, the voltage and current consumed by the motors. Lidar to precisely measure the altitude in relation to the water, and a PixHawk controller used to record accelerations, speeds, position, attitude of the aircraft, etc. As results are presented the energy consumption of the batteries as a function of altitude, in flight condition in ground effect, as well as the thrust force generated by the motors, in addition to determining up to which altitude that the ground effect has a good performance and improves the efficiency of energy consumption of the Volitan.

© 2024, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This work proposes an assessment of the delta wing sweep variation of a Generic Future Fighter in the conceptual design phase. Combat aircraft have critical control and therefore the stability analysis of these configurations is compared. Little variation in stability was observed between the 5 different configurations. This indicates that other requirements may become more relevant when designing a fighter aircraft, such as stealth and performance. Thus, this work aims to evaluate the impact of wing sweep on the longitudinal stability of fighter aircraft, considering five different sweep angles: 45°, 47°, 50°, 55°, and 60°. To conduct this analysis, a numerical evaluation, using the Vortex Lattice Method (VLM), wind tunnel results and parameter identification data from past work will be used to obtain the aerodynamic data for each configuration. The aerodynamic data will then be used in a time-domain flight simulation model to analyze the longitudinal stability of the aircraft.

© The Author(s), under exclusive license to Springer Nature Switzerland AG 2024.The aeroelastic typical section, also known as the three degrees of freedom (3DoF) aeroelastic model, is a common way to start studying aeroelastic systems, especially when there are nonlinearities that can be isolated. There is a lack of research using cubic springs controlling aileron deflection and considering Peters’ unsteady loading acting on the model simultaneously. The model presented here have two linear springs (one commanding the vertical displacement and the other commanding the pitch angle) and one nonlinear cubic spring for aileron deflection. Peters’ unsteady model is used to define the lift and aerodynamic moment, forces used in the flutter analysis. In addition, this model is validated for very flexible surfaces, such as helicopter blades. With the numerical simulated time series, the 0–1 test is performed, as well as the Takens reconstruction and the determination of the Lyapunov exponent. The 0–1 test result is compared to the Lyapunov exponent, as part of their validation for aeroelastic systems subjected to structural nonlinearities. With this validation, in future work, these methodologies shall be applied in a more complex aeroelastic system, which will be a flat plate clamped at the root.

© Published under licence by IOP Publishing Ltd.The motivation to accurately model the dynamics of flexible aircraft grew with the development of energy-efficient aircraft, consequently, great aspect ratio aircraft. The development of an accurate model that represents the flight dynamics of a flexible aircraft has been pursued by industry and aeronautical research organizations during the last decades. One of these approaches is to find a flexible aircraft model using systems identification methods. This research aims to apply an integrated model containing longitudinal and lateral directional rigid body dynamics, coupled to the first four flexible body modes, for identification and validation from flight test data. The Unmanned Aerial Vehicle (UAV) Eolo with the flexible wing is used during the experiments. Initially, a finite element structural model (FEM) based on beam elements, concentrated masses, and rigid bars was used. The quasi-stationary panel model based on the Vortex Lattice Method (VLM) was adopted for the aerodynamic model. Two diagonal matrices were used to correct the aerodynamic influence coefficients (AIC) matrix obtained via VLM before and post-multiplication for aircraft identification. The estimation of the main diagonal elements of each matrix was obtained through the Output Error Method in the time domain. A model validation analysis was carried out, which shows a good correlation between the model and measurement data. In conclusion, getting correction matrices instead of stability derivatives is beneficial because matrices can be used more directly during the aeronautical design and observe the behavior concerning loads.

© 2024, Academia Brasileira de Ciencias. All rights reserved.Multirotors Aerial Vehicles are special class of Unmanned Aerial Vehicles with many practical applications. The growing demand for this class of aircraft requires tools that speed up their development. Simulated environments have gained increasing importance, as they facilitate testing and prototyping solutions, where virtual environments allow real-time interaction with simulated models, with similar behavior to real systems. More recently, the use of Augmented Reality has allowed an increasing experience of immersion and integration between the virtual world and a real scenario. This work proposes the use of Augmented Reality technology and a simulated model of a multirotor to create an interactive flight environment, aiming to improve the user experience in the analysis of simulated models. For this purpose, a smartphone was adopted as a hardware platform, a game engine is used as a basis for the development of the Augmented Reality application, that represents a numerical simulation of the flight dynamics and the control system of a multirotor, and a game controller is adopted for user interaction. The resulting system demonstrates that Augmented Reality is a viable technology that can be used to increase the possibilities of evaluating simulated systems.

© 2024 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.Although in-flight refueling procedures are already widely performed thought military, most of all published guidelines and related documentation for certification/qualification between aircraft are focused on methods for Fixed-Wing Air to Air Refueling (FWAAR) receivers. Additionally, during an Air to Air Refueling certification process, cost efficiency is of utmost importance since it requires several aircraft during flight testing phase (tanker, receiver and usually a chase aircraft). Based on HAAR (Helicopter Air to Air Refueling) flight tests performed by the Brazilian Air Force (BAF) between Airbus H225M and Lockheed Martin KC-130H, this paper presents a statistical assessment tool that was developed in order to further investigate HAAR contact tasks results. A workload analysis was performed and compared to qualitative evaluations based on Cooper-Harper ratings for pilot input profiles on flight controls during contact tasks. Therefore, the main goal is to present lessons learned during HAAR flight tests as well as tools and methods that can be used to provide insight on which conditions should be further investigated, thus enhancing flight test efficiency.

© 2024 IEEE.This article aims to assist the evaluation of imaging models for deployment in public security helicopters, specifically for the Military Police of Rio de Janeiro State. It aims to establish a comprehensive technical framework for determining the crucial prerequisites of such equipment, employing a multicriteria decision support approach. In addition to indicating available solutions capable of ensuring the effective development of missions. The analysis combined two methods, the PSI (Preference Selection Index) and the CoCoSo, the Combined Compromise Solution, the first being used to determine the weights of the criteria and the second to evaluate the alternatives to each criterion, generating a solution that represents a compromise between the different options and based on this solution, classify them according to their overall adequacy. The evaluations of the models about the predefined criteria considered only the technical data provided by the equipment manufacturers without incorporating subjective criteria. The analysis of the results achieved through the methodological approach adopted constitutes a robust foundation to guide crucial decisions on the definition of the most appropriate imaging cameras for use in helicopters used in police air missions, as it clearly and objectively highlights the most advantageous options and aligned with the specific needs of this branch of activity. They are enabling decision-makers to make the right choices, which will play a significant role in improving the performance of security forces in their responsibilities, as well as providing a substantial increase in the protection of society.

© 2024 International Forum on Aeroelasticity and Structural Dynamics, IFASD 2024. All rights reserved.The objective of this work is to obtain CFD results for the dynamic response of a wing oscillating in pitch in a transonic regime using an open-source tool. The purpose is to verify and improve the correspondence with the experimental data as performed in the wind tunnel test for a wing model developed by Embraer. For this, in some analyzes it is proposed to impose a prescribed movement to the wing in the CFD simulations that models the bending observed in the scaled model throughout the tests as a rigid mesh movement in rolling direction. Prescribed motion parameters are extracted directly from the model’s structural deformation measurement data. In addition, simulations of a test case using the Benchmark Supercritical Wing (BSCW) are performed to investigate the impact of relevant variables in this type of analysis, such as time step and mesh refinement level. The time step was identified as the most influential parameter to approximate the simulation results to experimentally obtained data. The CFD results for the Embraer wing were able to capture the main behaviors of the magnitude and phase of the non-stationary pressure coefficient on the wing, mainly for conditions of higher reduced frequencies, with an affordable computational cost.

© 2023 by the authors.The objective of this article is to characterize an aeroelastic system in terms of its dynamical behavior, which could be either chaotic or periodic before, during, and after achieving the flutter velocity. The aeroelastic system shown here is a wing with a high aspect ratio, which leads to a very flexible behavior subjected to unsteady flow. This paper compares the computational and experimental dynamical behavior of an aeroelastic system at the flutter velocity for the different dynamic stall models proposed. To understand the nonlinear behavior of this system, the traditional attractor reconstruction and Lyapunov exponent calculation are compared with the 0–1 test. In addition to this comparison, two dynamic stall semi-empirical models are applied directly to the time history. All these comparisons show that the computational and wind tunnel experiments are in good agreement, and the dynamic behavior usually gives close results for the 0–1 test and Lyapunov exponent. It is concluded that the system presents chaotic behavior when no dynamic stall correction is applied or when Gangwani’s correction is applied. However, Boeing–Vertol’s correction postpones the chaotic behavior, meaning that the chaotic behavior is only observed for velocities above the flutter.

© 2023 IEEE.The paper describes the constitution of the Engineering Research Center for the Aerial Mobility of the Future (ERC-AMF) having ITA as the host institution, Embraer as the industrial partner, and researchers from the University of São Paulo and the University of Campinas. The objective of the ERC-AMF is the realization of R&D to contribute to overcoming challenges to the shaping of aerial mobility in the upcoming decades. These challenges arise from the necessity of reducing pollutant and noise emissions, and the need for increased efficiency of manufacturing processes, besides the trend of introducing in the market novel aircraft adapted for operation in urban environments and short-range travels. Five research areas are focused on the first operation phase of the Center: Machine Control for Electric Propulsion; Aeropropulsion Integration in Electric Aircraft; Methods for Decision Making in Autonomous Systems; Advanced Design for Metallic Additive Manufacturing; and Intelligent Aircraft Final Assembly. Each line will be developed by researchers from partner universities and engineers from Embraer. It is expected that the Center will contribute to the appropriation, by the Brazilian aeronautical industry, of scientific and technological knowledge generated, and, as a result, increase its preparedness to face challenges that shall be overcome in the process of shaping the aerial mobility of the upcoming decades.

© 2021 by Antônio B. Guimarães Neto, Guilherme C. Barbosa, Juliano A. Paulino, Rafael M. Bertolin, Jéssica S. M. Nunes, Pedro J. González, Flávio L. Cardoso-Ribeiro, Maurício A. V. Morales, Roberto G. A. da Silva, Flávio L. S. Bussamra, Flávio J. Silvestre, Fernando J. O. Moreira, and Carlos E. S. Cesnik. Published by the American Institute of Aeronautics and Astronautics,.The challenges of modeling flexible aircraft include appropriate fidelity capturing and validation with experimental data. In fact, the validation of formulations and models for the flexible flight dynamics is indispensable to ensure that all the important phenomena are correctly captured. With this objective, two high-aspect-ratio flexible aircraft have been flight-tested, and coupled aeroelastic–flight dynamics data have been collected to support model validation. Additional ground vibration and static tests were carried out to fully characterize the structural dynamic properties. Numerical models were built based on a linear structural representation but with geometrically nonlinear aerodynamics. Low Reynolds number effects were included in a simplified way with lookup tables of two-dimensional airfoil data. Wing-tip effects were considered via the vortex-and doublet-lattice methods. Propulsive data were obtained with wind-tunnel tests. This paper describes the numerical models, the two aircraft, and their instrumentation and presents the data collected from the aircraft sensors during flight tests. Numerical and experimental results are compared for angular velocities, accelerations, and strains measured at different points of the aircraft. Despite its limitations and simplifications, the numerical model captures the real aircraft main aeroelastic and flight dynamic behaviors.

© 2021, The Society for Experimental Mechanics, Inc.Piezoelectric materials have been increasingly applied to a wide range of engineering and scientific applications in the past three decades. One application of interest involves wind tunnel testing to quantify and evaluate the aeroelastic behavior of aircraft wings. In this paper, the focus is on the suitability of piezoelectric sensors, namely PVDF (Polyinylidene Fluoride), to quantify the aeroelastic response of wing models by acquiring modal parameters, i.e. the natural frequencies and damping factors of the vibration modes. Concurrently, a complementary goal is to use PZT (Lead Zirconate Titanate) materials as actuators to better excite the vibration modes that are not adequately energized by the aerodynamic flow. During the setup phase of the experimental apparatus, several studies were performed to help define the test parameters. The aeroelastic tests were conducted in a wind tunnel using a single PZT as actuator and a single PVDF as sensor. The modal parameters obtained using a single PVDF sensor response were then compared with those estimated using laser doppler vibrometry. These parameters were then used to estimate the pre-flutter speed, using both sensing techniques, for three case studies with different mass ballast configurations. A very good agreement was observed between the two sensing techniques, when comparing the results in terms of the frequencies and damping factors of the mode shapes leading to flutter. The results show the suitability of using a single PVDF sensor to estimate the modal parameters, in very turbulent and noisy conditions that are characteristic in wind tunnel testing. PZT is found to reduce the exogenous noise caused by the aerodynamic flow when considering a high number of averages.

© 2022, Emerald Publishing Limited.Purpose: While computational methods are prevalent in aircraft conceptual design, recent advances in mechatronics and manufacturing are lowering the cost of practical experiments. Focussing on a relatively simple property, the lift curve, this study aims to increase understanding of how basic aerodynamic characteristics of a complex stealth configuration can be estimated experimentally using low-cost equipment, rapid prototyping methods and remotely piloted aircraft. Design/methodology/approach: Lift curve estimates are obtained from a wind tunnel test of a three-dimensional-printed, 3.8%-scale model of a generic fighter and from flight testing a 14%-scale demonstrator using both a simple and a more advanced identification technique based on neural networks. These results are compared to a computational fluid dynamics study, a panel method and a straightforward, theoretical approach based on radical geometry simplifications. Findings: Besides a good agreement in the linear region, discrepancies at high angles of attack reveal the shortcomings of each method. The remotely piloted model manages to provide consistent results beyond the physical limitations of the wind tunnel although it seems limited by instrumentation capabilities and unmodelled thrust effects. Practical implications: Physical models can, even though low-cost experiments, expand the capabilities of other aerodynamic tools and contribute to reducing uncertainty when other estimations diverge. Originality/value: This study highlights the limitations of commonly used aerodynamic methods and shows how low-cost prototyping and testing can complement or validate other estimations in the early study of a complex configuration.

© 2022 by the authors. Licensee MDPI, Basel, Switzerland.In this paper, a dynamic scaling methodology is introduced to devise reduced scaled models of aircraft with the objectives of minimizing the development cost and exploring the design space. A promising way to accomplish this is using Topology Optimization (TO) for Additive Manufacturing (AM). Here, TO is employed to design a reduce scale model by matching its natural frequencies and mode shapes to those of a full scale model. Different TO strategies based on density approach are tested with the goal of achieving a dynamically scaled structure that can be manufactured. To achieve this goal, the TO solution should be free from intermediate densities, which is observed for some TO strategies but not all. When no penalization factor is applied: (i) the relative difference between natural frequencies is less than 1% and (ii) the estimated Modal Assurance Criteria (MAC) metric to evaluate the correlation between mode shapes is close to the ideal identity matrix. These results demonstrate the effectiveness of the dynamic scaling methodology. However, when using a penalization factor to avoid intermediate densities, the dynamic behavior correlation between full and scaled models degrades. This trend is more visible in the MAC metric, where off-diagonal terms above 20% and diagonal terms below 90% appear.

© 2022, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.With increasing pressure to lower pollutant emissions, the aerospace industry has turned its attention to the design of more efficient aircraft. Electric airplanes are seen as one of the most promising solutions to this problem, and significant investments are being made to develop this type of aircraft. Since the electric propulsion system is distinct from those based on internal combustion engines, the performance characteristics of all-electric airplanes can be significantly different from that of regular aircraft. An important element of this new type of propulsion system, and one of the reasons for its unique characteristics, is the power source. Fuel cells are one of the main embedded power sources employed to provide electricity in vehicles, and its use to power electric airplanes is currently being researched. This work presents an analytical investigation of fuel and oxidizer consumption during the cruise flight of all-electric aircraft powered by fuel cells. This study is relevant because cruise flight usually is the crucial phase that drives aircraft design requirements in what concerns energy requirements. A novel formulation is developed, and parametric models are provided for the airplane relevant systems. New analytical solutions are derived in parametric, closed form, allowing quick calculations and eliminating the need for numerical solvers and possible convergence issues. Also, simulations are provided to illustrate the method developed in the article. The results show that the optimal velocities for minimal consumption can be higher than predicted by conventional methods.

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.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.

© 2022, American Institute of Aeronautics and Astronautics Inc. All rights reserved.System identification based on mathematical models is generally restricted to linear systems. To model nonlinear behavior, more complex mathematical models are needed and often not available. To model the nonlinear dynamics at high angle of attack of a fighter, a neural network method was applied. The system identification process used in this work used flight test data acquired from a remotely piloted Generic Future Fighter (GFF) subscale. After the application of the neural network, the non-linear effect present in the detachment of the wing boundary layer was possible to estimate. The neural network used was the Feedforward type and the optimization of the parameters was carried out with Backpropagation. Stall maneuvers were initially used to train the neural network (training cycle) and later a new stall maneuver was used to validate the identification (prediction cycle). The method demonstrated the ability to estimate the lift curve in a subscale fighter.

© 2022, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.It is well-known that ice accretion can adversely impact the aerodynamic performance of airfoils and wings. In this work, we conducted an experimental investigation on the impact of different ice shapes on the flow around airfoils. The NACA 23012 and the GLC-305 airfoils were tested at a low-reynolds wind tunnel, which included forces, moments and surface pressure were evaluated, and Particle Image Velocimetry (PIV) was used for flow field measurement. The studied ice type was a simulated single horn based on the glaze ice accreted on airfoil leading edge, with different heights and chord position. The parametric approach was applied in order to vary the ice geometric characteristics. Evaluation was performed with the ice shape extruded throughout the entire span of the airfoil, and the objective of this research was to provide a flowfield-physics perspective on the flow with different ice geometries and its effect on the overall aerodynamic performance of the airfoil under low Reynolds conditions.

© 2022, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Although there are literature references that detail guidelines and flight test techniques for Fixed Wing Air-to-Air Refueling (FWAAR), there are no well established methods to perform Helicopter Air-to-Air Refueling (HAAR). Straightforward application of those FWAAR methods neglecting specific helicopter performance and HQ characteristics did not demonstrate to be a successful approach, given that actual critical conditions for defining HAAR envelope could not be identified. Therefore, this work presents methods and techniques developed by the Brazilian Flight Test and Research Institute (IPEV) during the HAAR qualification process between the helicopter H225M and the tanker KC-130H. Results demonstrated that a more complete assessment for defining HAAR envelope could be performed when using a power margin approach for planning and performing flight tests. Literature CHR (Cooper-Harper Rating Scale) contact tasks for FWAAR based on precision performance criteria were then tailored in order to take into account low closure rate profiles and limited power margins. Additionally, given the limitations of the CHR assessment, a modified version of DIPES (Deck Interface Pilot Effort Scale), which is well-know for multi-axis evaluation in helicopter / ship qualification flight testing, was also applied when performing contact tasks. This approach allowed to identify unacceptable pilot workload levels more easily than the CHR. Therefore, this paper aims at presenting the lessons learned during planning, execution and data processing steps of the HAAR flight test campaigns in order to further enhance flight test techniques on that type of procedure.

© 2021 The Authors. Human Brain Mapping published by Wiley Periodicals LLC.Understanding decision-making in complex and dynamic environments is relevant for designing strategies targeting safety improvements and error rate reductions. However, studies evaluating brain dynamics in realistic situations are scarce in the literature. Given the evidence that specific microstates may be associated with perception and attention, in this work we explored for the first time the application of the microstate model in an ecological, dynamic and complex scenario. More specifically, we evaluated elite helicopter pilots during engine-failure missions in the vicinity of the so called “dead man's curve,” which establishes the operational limits for a safe landing after the execution of a recovery maneuver (autorotation). Pilots from the Brazilian Air Force flew a AS-350 helicopter in a certified aerodrome and physiological sensor data were synchronized with the aircraft's flight test instrumentation. We assessed these neural correlates during maneuver execution, by comparing their modulations and source reconstructed activity with baseline epochs before and after flights. We show that the topographies of our microstate templates with 4, 5, and 6 classes resemble the literature, and that a distinct modulation characterizes decision-making intervals. Moreover, the source reconstruction result points to a differential activity in the medial prefrontal cortex, which is associated to emotional regulation circuits in the brain. Our results suggest that microstates are promising neural correlates to evaluate realistic situations, even in a challenging and intrinsically noisy environment. Furthermore, it strengthens their usage and expands their application for studying cognition under more realistic conditions.

© 2021 IEEE.This paper presents a method to synchronize data acquisition devices that are mechanically coupled, having attached an accelerometer to each device. A common time base for the accelerometer signals are obtained through the identification of pairing salient peaks and applying line-fitting through the potential matches. Aligning data recorded from different sources is important to precisely provide an observation of the state of a system in time (sensor fusion), to estimate the correlation between its variables and to correlate variables to time-based events. A data link between devices is not always possible or convenient. If the acquisition devices are mechanically coupled, such as being in the same body or vehicle, we propose to synchronize the data recorded from both by using the accelerometers signals to bridge. The provided solution is an automated process to find the temporal reference between accelerometer signals. Several signal processing steps are taken after data collection and storage: inconsistency removal and filtering, detection of maxima and minima, selection of saliencies, description through a characteristic pair of numbers: the interval lengths between it and its successor and its predecessor, listing possible matches between salient points, selection of the topmost relevant matches and line fitting with consensus. We discuss qualitative similarities of related work. Quantitative results are also presented by using the multidisciplinary study that motivated this work, with simultaneous data from the instrumentation of a helicopter and pilot physiological data. To conclude, we discuss the limitations of the presented approach and future work.

© 2020 IOP Publishing Ltd Printed in the UKTemperature-sensitive paint (TSP) is a type of luminescent temperature sensor that can provide surface temperature measurements on the surface of a test article. These surface temperature measurements can be used to determine heat flux and visualize boundary layer transition on an aerodynamic body. This paper presents the static response characteristics of a TSP based on rhodamine B (RhB) applied on a polymer-ceramic supporting matrix. Six solvents with varying polarity indices were used to apply the RhB to the sensors, which were then evaluated in terms of luminescent signal level and temperature sensitivity. The results are presented and also compared to previous results for RhB on an anodized aluminum matrix. The temperature range was between 150 K and 365 K. The study confirms that the influence of the solvent on the sensor's final characteristics is significant, and shows temperature sensitivities as high as −4.8% K−1 at 150 K when dichloromethane is used as the application solvent.

© 2020, The Society for Experimental Mechanics, Inc.The new aeronautical structures have become more flexible and light weight to reduce energy consumption, emissions, and noise, and to operate at high altitudes for long periods in the air, such as those required in the NASA Helios project. The increased structural flexibility of these aircraft has reignited concerns related to aeroelastic instabilities, such as flutter. Improving the techniques and methods used in aircraft certification flights is an important concern of the aeronautical community, because current standards and procedures do not provide recommendations and guidelines for aircraft with a high degree of flexibility. The techniques traditionally used in commercial airplanes cannot be used in this new aircraft concept, because they have a high degree of non-linearity in their flight dynamics. Current research indicates an increasing awareness about the importance of vision in the monitoring of UAV structural health. This work presents a new methodology to measure natural frequencies of aeronautical structures using a computer vision system. We also discusses new approaches to sense and acquire vibration data on aeroelastic certification flights test. These new approaches aim to reduce both the time required to identify the aeroelastic phenomenon and the size of the hardware that must be boarded on the aircraft, thus minimizing the risks and costs of the vibration tests. The advance of computer vision systems enables the use of cameras as a motion tracker sensor with millimeter precision and accuracy. Non-contact sensors are suitable for flutter analysis because they do not interfere with the dynamics of the aircraft. Therefore, this new methodology is able to process the obtained images and provide the user with the data about movements in a ready to use vector, at a reasonable cost. Using the data provided by this methodology the natural frequencies of the first bending modes were identified. This new methodology can be a user-friendly tool to support the Brazilian National Civil Aviation Agency - ANAC program called iBR 2020, which aims to certify small aircraft.

© 2020, Krishtel eMaging Solutions Private Limited.purpose: This paper aims to investigate the post-flutter nonlinearities due to the wing high aspect ratio. Flutter is an aerodynamic auto-excited phenomenon which occurs due to the coupling of two or more different vibration modes. This coupling results from the interaction between aerodynamic, elastic, and inertial forces. The present investigation scope is the comparison between experimental aeroelastic analysis of high-aspect-ratio wings using subsonic wind tunnels with computational experiments for three different center of gravity conditions. Methods: The computational nonlinear aeroelastic analysis is performed using a formulation based on the variation of the energy functional and considering Peters’ unsteady aerodynamic model. The post-flutter analysis consists in the comparison between the attractor reconstructed using the Takens’ theory with the result of the 0–1 test for chaos for both computational and wind tunnel experiments. Results: First, the 0–1 test is performed for both computational nonlinear analysis and wind tunnel experiment and they are in good agreement for each condition. After that, the procedure to obtain the reconstructed attractor is followed for both cases and they are again in good agreement. The difference between computational experiment reconstructed attractors and wind tunnel experiment reconstructed attractors is mainly because of wind turbulence in wind tunnel section. Conclusion: This research shows that all three experimented conditions presented periodic dynamic behavior. The 0–1 test predicted well the dynamic behavior; therefore, it is recommended to apply this test for all nonlinear experiments so that a first qualitative evaluation is possible. All Lyapunov exponents calculated are negative, confirming the periodicity of the three center of gravity conditions for both computational and wind tunnel experiments.

© 2021 32nd Congress of the International Council of the Aeronautical Sciences, ICAS 2021. All rights reserved.Aiming to build a 6DOF flight simulation model for NASA-CRM and, in the lack of experimental lateral-directional data, three different methodologies were assessed and compared: usual semi-empirical method based on historical wind tunnel experiments and first principals physics, vortex lattice method (using XFLR5 software), and panel method (using OpenVSP software). The comparison presented is exhaustive in the aspect of isolating each component aerodynamic contribution and presents a clear view of the agreement between the studied methodologies. This work serves as basis for a first assessment of the uncertainty of the stability derivatives, associated with each methodology, usually used during preliminary design of aircraft.

© 2021 32nd Congress of the International Council of the Aeronautical Sciences, ICAS 2021. All rights reserved.A method for calculation of dynamic response to gust and control surface excitations, which allows the inclusion of nonlinear terms such as nonlinear control laws, without the need to use rational function approximation, is presented. This method is based on characteristic responses obtained with frequency domain equations, which associated with Fourier transforms, can be used to calculate time domain dynamic responses via Duhamel Integrals. The Duhamel Integral method is demonstrated for a typical airfoil section with three degrees of freedom, making use of frequency domain equations and direct and inverse Fourier Transforms for the characteristic response computation. In addition, applications where this method presents advantages, such as in the application of non-linear control laws and analysis of Oscillatory Malfunction (OMF), are demonstrated.

© 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, 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 LtdCurrent trends in the aircraft industry involve higher aspect-ratio wings made of lighter materials. These trends seek to reduce fuel emissions and increase flight efficiency by reducing drag to lift ratio and overall weight, respectively, of the aircraft. This results in reduced structural stiffness and coupling between the aeroelastic modes and flight dynamics. The flutter phenomenon is of particular interest for aeroelastic studies, and modeling post-flutter limit-cycle oscillations (LCO) is a challenging problem. Several studies have been developed to allow fast simulations of the highly non-linear aerodynamic situations, with leading-edge vortex modulation been a proved solution for modeling some forms of LCOs in airfoils. This article proposes a framework based on the 3D expansion of this method using strip theory and coupling with modal structural model for simulations of aerodynamic based non-linear phenomenon. A cantilevered flat plate is used for testing and validating the framework against wind-tunnel experiments and the industry standard approach. The results show that the proposed model is able to capture the main behavior of the LCO observed in the experiments and is directly comparable with the current approaches used at the industry. The framework allows for scalability and is also fast enough to provide time-based results in under two days for a desktop simulation, reducing the need of expensive cluster computations. Finally, since it is completely physics-based it allows for the engineer to get insights on the aerodynamic flow at a fraction of the cost of more detailed CFD models.

© The Institution of Engineering and Technology 2020.This chapter discusses new approaches to sense and acquire vibration data and to pre-process these data on aeroelastic certification test flights. These new approaches aim to reduce the time to identify the aeroelastic phenomenon and to reduce the size of hardware that must be boarded in the aircraft, thus minimising the risks and costs of the vibration tests. The presented experiments construct a way to develop a non -contact measurement system for flight vibration tests in the aircraft certification process. These experiments have shown that the techniques used today for in-flight trials will be obsolete in the near future, as the aeronautical structures are becoming lighter every day, thus not admitting any additional mass for instrumentation in-flight trials.

© 2020 American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Electric propulsion has become a subject of great interest in the aeronautical community. The present work focuses on the optimal climb performance of such aircraft, in the sense of minimizing the electric charge consumed. This analysis is relevant, since this flight phase is very energy consuming and demanding on the propulsion system. A cost function – the electric charge consumed – is derived in parametric form, with both aircraft and motor parameters. Analytical expressions for the optimal lift coefficient, throttle, velocity and electric charge are derived as solutions for a static optimal control problem. The solutions are also presented in parametric form, being useful for applications in conceptual design and performance analysis. This is supplemented by simulations in order to provide a quantitative insight to the problem.

© 2019, The Brazilian Society of Mechanical Sciences and Engineering.The interest in applying piezoelectric materials for modal analysis has been growing in the past few decades. In piezoelectric materials, both electrical and mechanical domains are coupled, i.e., these materials are able to convert electrical energy into mechanical energy and vice versa. Due to this key characteristic, they can be used in several applications as actuators or sensors. Furthermore, some piezoelectric materials exhibit a predominant coupling, which makes them more efficient when used for specific purposes/applications. This is the case of the polyvinylidene fluoride (PVDF) which is widely used as a sensor. An advantage associated with the PVDF is its small influence on the results, due to the low thickness and high flexibility; sometimes, its influence is completely neglected. The aim of this work is to evaluate the influence of a single PVDF film on a flexible beam model. For this purpose, an efficient methodology to verify and identify the intrusiveness level of the instrumentation is proposed, which consists in changing the sensor position (PVDF) and simultaneously acquiring the data by using a non-intrusive technique (laser vibrometer). The modal parameters (natural frequencies and damping factors) obtained by PVDF and laser vibrometer responses should be very close for each PVDF position. If this condition is satisfied, the variation of the modal parameters due to PVDF position will show the intrusiveness level imposed by the PVDF instrumentation. This research emphasizes the importance of verifying the influence of the instrumentation, even if it seems to cause merely a small intrusiveness on the dynamic system.

© 2019 Elsevier Masson SASAmong the aeroelastic phenomena most commonly affecting flexible and very flexible aircraft, those caused by gusts deserve special attention due to their potential either in degrading flying qualities and ride comfort or in increasing structural loads. It is then of interest to structural loads and flight controls engineers that solutions be developed to attenuate the effects of gusts on aircraft. Particularly, the use of piezoelectric transducers arises as one of the potential solutions in the design of gust load alleviation and structural mode suppression systems. In this paper, the gust load alleviation on a flexible smart idealized wing using only piezoelectric transducers is analyzed and experimentally tested. The numerical model includes a finite-element model of the wing, employing two-node, seven-degree-of-freedom smart beam elements, assuming small deformations and neglecting transverse shear. A quasi-steady, strip-theory-based aerodynamic model is used. Two control laws are evaluated: one based on output feedback, and the other based on feedback of observed states of a truncated system. Using a gust generator, wind-tunnel tests were performed at different flow speeds and gust frequencies to validate the computational model and to verify the performance of piezoelectric transducers. The results show a considerable attenuation of the wing root bending moment, especially using two piezoelectric actuators. Important performance improvements were overall verified with feedback of observed states when compared with static output feedback, specially to decrease the participation of the elastic modes in the gust response.

© Statement The authors confirm that they, and/or their company or organization, hold copyright on all of the original material included in this paper.The workload assessment to perform a full autorotation on the AS-350 aircraft (Airbus Helicopters) was performed during a Flight Test Campaign with 80 flight hours and 227 data collection procedures, considering 10 pilots with different piloting skill levels, among such military pilots, flight instructors, and test pilots. During the tests, these pilots were subjected to unexpected engine failures, to evaluate the actual reaction time of each pilot, and to test the ability to make a safe landing under the conditions prescribed by the aircraft manufacturer. The testing method used began with unexpected engine failures when only the lead test pilot knew that the engine failure would be simulated. In the sequence, several points of autorotation were performed, from the simplest profile to the most complex. All the procedures have registered the performance parameters and handling qualities of the aircraft, along with the physiological parameters of the pilots. The aircraft was equipped with dedicated instrumentation for in-flight testing and the pilots have been instrumented with an Electroencephalogram (EEG), Electrocardiogram (EKG), Respiration Belt and Galvanic Skin Response (GSR), Eye Tracking and Face Recognition Camera equipment. This instrumentation was employed to determine physiological markers that could determine the pilot workload, quantitatively, reducing the subjectivity of measures that use only qualitative scales of evaluation, such as Handling Qualities Rate (HQR) and Bedford Workload Scale (WL). In this work, only the preliminary results of the analysis obtained by the Galvanic Skin Response markers will be presented. Major potential applications of the results from the present research range from cockpit design guidelines and human-machine interface systems for supporting pilotage such as more effective alarm systems, interactive cockpits, enhancement of active autopilots with semi-automatic flight commands. Besides that, the results and conclusions from this research can also improve processes and methods for the training-based formation of pilots, along with the development of flight simulators with physiological measurements parameters quantification, feeding back data for a piloting performance assessment.

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Aggressive targets for reducing aviation environmental impacts require considerable improvements in aircraft design such as nonconventional aerodynamic configurations, new materials and more efficient systems. This scenario poses airframe manufacturers to make more efficient products only achievable by making a complete integrated design between airframe and systems. In this sense, the aircraft design shall be committed to the maximum efficiency and minimum waste of useful energy (exergy destruction). The concept of exergy analysis has already been successfully applied to evaluate, compare and optimize thermal systems and chemical processes in other industrial fields. Thereby, this paper presents how exergy analysis can be used as a tool for aircraft design and shows an application of the method to evaluate competitive propulsion system architectures for a regional aircraft.

© Universal Technology Corporation, 2018.The design of the conventional configuration of commercial aircraft, composed by a tube fuselage, a cantilever wing and an empennage, has been improved since its introduction in the 1950s and it is unlikely that great improvements should occur without drastic changes. The strut braced wing aircraft presents itself as an option. The main difference lies on a strut connecting the wing to the fuselage, reducing the bending moment of the wing and, consequently, its weight. Alternatively, the wing span could be increased, or even the wing thickness decreased, without great weight penalties. This combination of geometric changes reduces drag and improves performance. To evaluate possible aeroelastic issues that might hinder the development of this configuration, a parametric flutter analysis is performed based on aircraft of regional aviation size. As a result it was observed that: (a) increasing the wing aspect ratio from 8.3 to 12 decreases the flutter speed in 20%; (b) if the engine is positioned exactly at the wing and strut intersection at 70% of the wing span instead of 50%, an increase of 30% in flutter speed is obtained; (c) and that the flutter speed can be increased by 35% if the spanwise wing and strut intersection is moved from 70% to 30 % of the span.

© Universal Technology Corporation, 2018.Nonlinear aeroelastic phenomena are continuously investigated in aeronautical researches. The nonlinearity nature can be aerodynamic or structural. This work will investigate aeroelastic nonlinearities in a very flexible wing. A flutter analysis is proceeded in order to evaluate the error between the computational results and the experiment. Since the linear flutter theory considers small disturbances, nonlinear phenomena are expected. Both wind tunnel and computational experiments time series shall be analyzed and the evaluation if the system presents chaotic behavior will be performed through the 0-1 test.

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This paper identifies, characterizes and describes correction procedures for uncertainty sources associated with fast-response pressure sensitive paint (PSP) optical and image acquisition systems, including fixed pattern noise, non-linear CMOS sensor response and optical vignetting. Data on a typical CMOS camera was acquired using an easy to manufacture integrating sphere, and image correction procedures were developed having a typical wind tunnel environment in mind. For validation, corrections were applied to data on pressure distribution over a NACA 0012 airfoil in transonic flow. A significant improvement in data accuracy was observed.

© 2019 by German Aerospace Center (DLR). Published by the American Institute of Aeronautics and Astronautics, Inc.Aircraft equipped with an electric propulsion system are of increasing interest to the aeronautical community. This work focuses on the optimal performance of such aircraft during the climb phase, since it can be very energy consuming and demanding to the propulsion system. Analytical expressions for the optimal lift coefficient, throttle and velocity are derived as solutions to optimal control problems, for steady climb and for a simplified, linear accelerated climb, without a specified final time. The total energy cost and altitude gain are also derived in parametric form, allowing quick, useful estimations for conceptual design and performance analysis. This is supplemented by simulations in order to provide a quantitative insight to the problem.

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Flow and noise predictions of the tandem cylinder benchmark are performed using the open-source computational fluid dynamics code (CFD) Stanford University Unstructured (SU2). The numerical results were obtained using the Delayed Detached Eddy Simulation (DDES) approach with the new shear-layer adapted sub-grid length scale (SGS) for faster transition between RANS and LES. The Ffowcs Williams–Hawkings (FWH) analogy is used to propagate the pressure fluctuations to the farfield. Both flow and noise results are compared with experimental measurements from the Basic Aerodynamic Research Tunnel (BART) and Quiet Flow Facility (QFF) at NASA Langley Research Center. The compressible simulations are carried out on the mandatory 2-D grid from the ATAAC project with a spanwise extension of 3.0D, two different spanwise discretization were used to analyze the discretization effect. The comparison shows that the shear-layer adapted (∆SL A) is less grid dependent and more accurate than the standard (∆max) SGS.

© 2017, The Author(s) 2017.The main concern related to the flutter phenomenon is predicting and avoiding it. This paper describes the application of a flexural-torsional flutter testbed for acceleration reduction by applying active and passive model-based control. The model consists of the 2D typical section, with aerodynamic loads estimated by an unsteady time-domain formulation based on Wagner’s function. The active control architecture consists of a stability augmentation system with output feedback and gain scheduling via the linear-quadratic regulator theory and actuation by servomechanism. The passive control employs a shape-memory alloy to provide additional torsional stiffness. Experimental results show considerable reduction of oscillations at a relative low cost for both active and passive control strategies, and that the use of shape memory alloys in aeroelastic stability problems is promising.

© 2018, Springer-Verlag GmbH Germany, part of Springer Nature.We study a modified backwards-facing step flow, with the addition of two different plates; one is a baseline, impermeable plate and the second a perforated one. An experimental investigation is carried out for a turbulent reattaching shear layer downstream of the two plates. The proposed setup is a model configuration to study how the plate characteristics affect the separated shear layer and how turbulent kinetic energies and large-scale coherent structures are modified. Measurements show that the perforated plate changes the mean flow field, mostly by reducing the intensity of reverse flow close to the bottom wall. Disturbance amplitudes are significantly reduced up to five step heights downstream of the trailing edge of the plate, more specifically in the recirculation region. A loudspeaker is then used to introduce phase-locked, low-amplitude perturbations upstream of the plates, and phase-averaged measurements allow a quantitative study of large-scale structures in the shear-layer. The evolution of such coherent structures is evaluated in light of linear stability theory, comparing the eigenfunction of the Kelvin–Helmholtz mode to the experimental results. We observe a close match of linear-stability eigenfunctions with phase-averaged amplitudes for the two tested Strouhal numbers. The perforated plate is found to reduce the amplitude of the Kelvin–Helmholtz coherent structures in comparison to the baseline, impermeable plate, a behavior consistent with the predicted amplification trends from linear stability.

© 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.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Common aircraft mainframes are traditionally designed with unswept or backward swept wings. However, forward swept wings even presenting interesting aerodynamic characteristics were rulled out of aircraft design due the structural design challenges related with the torsion-bending coupling in static aeroelastic divergence. This paper proposes an investigation of sweep angle effect on aeroelastic response of wings made of isotropic and/or fiber reinforced composite materials with constant and variable stiffness. The aeroelastic behavior of flexible isotropic swept wing obtained with the implementation of a numeric aeroelastic system with finite elements and panels methods, for structure and aerodynamic, respectively, is compared with literature and validated with wind tunnel tests. The potential of tailored tow-steered laminated to enhance aeroelastic response is presented and compared with traditional unidirectional fiber reinforced carbon fiber-epoxy resin composites.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.An evaluation of the transonic buffeting onset boundary estimated by trailing edge pressure and RMS(root mean square) data of wing root strain is presented. The analysis is based on wind tunnel data of two modern transonic aircraft. The results obtained with the pressure and wing vibration data are compared to a steady aerodynamics coefficient methodology that was evaluated with flight test results. The evaluation concludes that strain gauge data provides a mean of quantifying the buffeting onset magnitude that can be correlated to flight test results. The trailing edge pressure data divergence criterion underestimates the buffeting onset boundary, but the magnitude of the pressure divergence can be adjusted for better results.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.The search to increase aircraft performance has led to wings with higher aspect ratios. High aspect ratio wings are subjected to aeroelastic instabilities. Consequently, the use of advanced materials added to its structure might be an aeroelastic control strategy. Among the passive and active methods to control and mitigate such structural phenomena, the use of shape memory alloys (SMA) has gained space in aeronautical applications. In such context, the present research aims to analyze the flutter answer of a flexible wing with high aspect ratio with SMA wires as passive controller. First, a research model has been designed, built and tested in a wind tunnel. After being calibrated, SMA wires were installed as a dynamic actuator in such model. The flutter condition was then evaluated through monitoring aeroelastic damping effect and the frequency coupling. The results showed lower peaks of the frequency response function (FRF) for some wire configurations as their critical flutter airspeed was reached. In addition, similar reductions of the limit cycle oscillations (LCO) were also observed.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.In order to understand the mechanisms behind self-sustained shockwave oscillations under laminar flow, the pressure distributions over a NACA 0012 airfoil with free transition were measured using two types of fast-response pressure-sensitive paints. The airfoil was submitted to flow near and at buffeting conditions. Pressure fields were analyzed using Power Spectral Density (PSD), Cross Power Spectral Density (CPSD) and pressure time-series standard deviation, as well as general flow visualization. Results indicate a clear shockwave oscillation frequency which does not match its turbulent flow counterpart. In addition, CPSD phase shift analysis allows the identification of different regions of shockwave interaction.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.This paper evaluates a mathematical model of 3 degrees of freedom (DOF), with different coupling factors between flaps (leading and trailing edges), to increase flutter velocity, known as the LAMBIE model [1]. This passive control airfoil concept with mechanical coupling factor between flaps, decreases camber when loading increases and increases camber when loads are smaller. To determine the structural matrices, the Lagrange equation is used and the aerodynamic matrices are determined by the Theodorsen model. This study contemplates simulation results of aeroelastic stability in the frequency domain. The method employed for the calculation of flutter is V-g [4]. In these simulations performed in MATLAB® software, it is possible to increase the velocity at which the flutter phenomenon occurs in 24.41%.

© 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.

© CSP - Cambridge, UK; I & S - Florida, USA, 2018.The aeronautical industry is continuously investigating nonlinear phenomena that might happen as it evaluates. Every dynamic system is subject to nonlinear behavior, especially if it is very complex like an aircraft. The nonlinearity nature, for these cases, can be aerodynamic, such as dynamic stall or shock waves, or structural, for example, freeplay or large displacements and high flexibility. This work will investigate a very flexible wing with high aspect ratio subjected to unsteady flow with a slender body at the wing tip to induce flutter. A flutter analysis is proceeded in order to evaluate the error between the computational results and the experiment. Since the linear flutter theory considers small displacements, nonlinear phenomena are expected. So the experiment time series shall be analyzed and this nonlinearity studied. The evaluation if the system presents chaotic behavior will be performed through the 0-1 test.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This work investigates wavy leading edge phenomena at transonic flow regime. Pressure sensitive paint measurements are performed over upper surface airfoil at transonic flow regime. The experimental investigation was conducted at a transonic wind tunnel at Reynolds number of 1,000,000, Mach number from 0.6 to 0.7, and angle of attack from 0 to 4 degrees. Two sets of airfoil were used, a smooth NACA0012 profile as baseline model and a wavy leading edge NACA0012 profile with amplitude of 3% and wavelength of 11% both related to chord of the airfoil. The models were manufactured using a 3D rapid prototyping which significantly improved time and cost, and also the feasibility and accuracy of such complex wavy leading edge airfoil. Pressure sensitive paint measurements indicates an impressive modification on flow pattern caused by tubercles when compared to baseline airfoil. If on hand, the baseline airfoil presents lambda-shock wave pattern, on the other hand, the wavy leading edge model changes this flow pattern avoiding shock wave structure. A likely explanation for tubercles avoid shock wave is related to possible counter-rotating vortex generated by wavy configuration upstream of the shock wave line. Thus, the results presented here indicates a potential to apply tubercles in commercial aircraft wings at transonic regime in order to decrease drag rise.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.The study of wavy leading edge phenomena on finite wings conducted on recent research has shown results that demonstrate improvements in aerodynamic characteristics for certain configurations. The most important outcomes were observed on swept and swept-tapered wings, attaining improvement in lift coefficients of around 20% when compared with their equivalent baseline models. This increase on lift force is associated to the fact that the wavy configurations exhibit a delay on the stall angle due to the effect of the tubercles, which main effect is to delay the stall progression from tip to root by keeping the flow attached on the leading edge at high angles of attack. Visualization results confirmed larger effect of these phenomena on the wingtip area. In order to give continuity to previous works that investigated swept wing with wavy leading edge, obtain a deeper knowledge of this phenomena and delimit the design space in which the wavy leading edge could be efficiently applied, a series of experiments were conducted on sixteen wing configurations including swept angles of 30 and 50 degrees, taper ratio of 1 and 0.5, and wavy span length of 20, 40 and 100% (from tip to root). All models had an underlying NACA 0020 airfoil and a wavy geometry with amplitude A = 0.03 and wavelength λ = 0.11 considering the root chord as reference. The purpose of this research consist in evaluating of drag and lift forces at Reynolds number of Re = 200, 000 for all models showing comparative results with the baseline wings. In addition, a flow visualization analysis using oil technique was included in order to better understand the involved phenomena.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.As a passive flow control mechanism inspired in nature, the wavy leading edge modifications have been tested here for specific rotary-wing airfoils. After examining previously studies, it was observed that these geometric devices could delay the boundary layer separation that usually occurs at high angles of attack in retreating blades, avoiding their abrupt stall. In order to evaluate if wavy leading edges could be applied successfully on helicopter blades, a series of wind tunnel tests have been performed for selected rotary-wing airfoils used in the H-60 Black Hawk aircraft. The waviness has shown a great potential to make softer the stall characteristics for the Sikorsky SC1094-R8 airfoil, a modified profile of the Sikorsky SC1095, while delaying the initial stall by up to 3◦, without a relevant increase in drag coefficient. The experimental investigation was based on force measurements (lift and drag coeficients) and oil flow visualization.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.To asses the state of art and best practices in computational aeroelasticity (CAe) for static and dynamic phenomena is the key objective of the AIAA Aeroelastic Prediction Workshop series (AePW). The workshop is an excellent opportunity for academia and in- dustrial exchange through transparent discussions and collaborative learning on relevant aeroelastic topics. During the second edition of the event (AePW-2), held in January 2016, the efforts were concentrate on unsteady aerodynamics and utter prediction for the Benchmark Supercritical Wing (BSCW). Three transonic cases were proposed: steady and unsteady forced aerodynamics under attached flow; utter on set prediction for weak shock attached flow condition; and a third more challenging case of transonic detached flow with steady, forced pitch oscillations and utter on set prediction. The workshop discussions pointed to some relevant issues in CAe while analyzing the participants results: turbulence modeling, temporal convergence, mesh convergence and fluid-structural coupling effects. Each analysis team explored some of those aspects, but not a consensus was established as best practices for numerical setup. As Computational Fluid Dynamics (CFD) become massive parallel processed, the number of grid points applied to relatively simple geometry as the BSCW becomes millions quickly. Parametric numerical studies of unsteady aero- dynamics and aeroelasticity including viscous effects in such meshes become too expensive for most researchers or industries and the physics of fluid flow and dynamics analysis can go to second plan faced the computational efforts to run and post-process such amount of generated data, risking to loose the engineering feeling of the analysis. This paper proposes the inclusion of a verification study case for the upcoming AePW editions in a lightweight 2D configuration, where some of the questions raised during the AePW-2 could be parametrically clarified regarding the CFD turbulence modeling, fluid-structure coupling and time and grid convergence. The aim is to explore the flow physics and see how the numerical setup behaves, identifying the limitations of the CFD methodologies applied, before jump into a fully 3D buffet aeroelastic configuration. The first results of this initiative are presented here as a CFD characterization of the BSCW airfoil section in steady, unsteady, unsteady forced and utter cases for the transonic regime of interest. All the data, models details and meshes are made available for the research community.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.We study a modified backwards-facing step flow, with the addition of two different splitter plates; one is a baseline, impermeable plate and the second a perforated one. An experimental investigation is carried out for a turbulent reattaching shear layer downstream of the two plates. The proposed setup is a model configuration to study how the plate characteristics affect the separated shear layer, and also how turbulent kinetic energies and large-scale coherent structures are modified. Hot-wire measurements show that the perforated plate changes the mean profile, mostly by reducing the intensity of backflow close to the bottom wall. Disturbance amplitudes are significantly reduced up to 5 step-heights downstream the trailing edge of the plate, more specifically in the recirculation region. A loudspeaker is then used to introduce phase-locked, low-amplitude perturbations up- stream of the splitter plates, and phase averaged measurements allow a quantitative study of large-scale structures in the reattaching shear-layer. The evolution of such coherent structures are evaluated in light of linear stability theory, comparing the eigenfunction of the Kelvin-Helmholtz mode to the experimental results. We observe a close match of linear- stability eigenfunctions with phase-averaged amplitudes for all tested Strouhal numbers. The perforated plate is found to reduce the amplitude of the Kelvin-Helmholtz coherent structures in comparison to the baseline, impermeable plate, a behavior consistent with the predicted amplification trends from linear stability.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This paper presents a detailed overview of hybrid RANS/LES methods as implemented within the open-source SU2 software package. We focus on the extensions of the existing RANS framework based upon the Spalart-Allmaras turbulence model that are necessary to apply the Delayed Detached-Eddy Simulation (DDES) technique. Particular emphasis is placed upon the low dissipation and low Mach number convective schemes required to maintain accuracy within the context of performing DDES in a second-order, finite volume, unstructured flow solver. We conclude with a suite of test cases across different regimes to demonstrate our DDES capability on both academic and industrial-grade applications.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.A new criterion, based on the chordwise movement of the aerodynamic center, is proposed to estimate the transonic buffeting onset of transport aircraft. Wind tunnel results obtained on four different aircraft are used to evaluate three existing traditional criteria that are based on CL x α and CM x α curves. The wind tunnel tests were performed at a chord Reynolds number of 3 million and the results were extrapolated to flight Reynolds number using cryogenic wind tunnel data for two similar aircraft. The proposed criterion presents the best predictions when compared to flight test data.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.A passive self-adaptive camber airfoil proposes to reduce fatigue and increase durability of lifting surfaces lifetime. The airfoil stability is investigated in an incompressible flow using a two-dimensional four degree of freedom model on pitch, plunge, slat and flap angles, in which the kinematic of slat and flap are coupled by a linear relationship. Concentrated cubic structural restoring forces as effect of nonlinearities are considered. The aeroelastic governing equations are written and integrated numerically using a fourth order Runge-Kutta scheme for the time domain evaluations. The identification and stability analysis of limit cycle oscillations are evaluated in the time domain by the Duhamel formulation and compared to the quasi-steady approximation and a method in the frequency domain using describing functions combined with the Sherman-Morrison formula. Investigations revealed that system parameters and initial conditions are crucial for the system dynamic stability that could lead to basin of attractions of periodic motions, heteroclinic orbits, jump phenomena and chaos.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This paper reports the current status of a joint Swedish-Brazilian research project aiming at exploring sub scale flight testing. A 13% scale fighter aircraft is used as a test bench for developing methods and procedures for data acquisition. This paper will present an Aerodynamic database as a partial result of the collaborative project.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All Rights Reserved.Futuristic aircraft designs and novel aircraft such as High Altitude Long Endurance (HALE) involve a higher level of structural flexibility than in conventional aircraft. Even at present, the trends in the aviation industry are to increase wing length (to reduce induced drag) and maximize use of composites, which lead to increased structural flexibility. This necessitates a rethink of conventional (linear) aeroelastic analysis, since the increased flexibility results in coupling between the flight dynamic and aeroelastic dynamics, and consequently, limit-cycle oscillations of the structure. In this paper, a new three-dimensional low-order model for unsteady aerodynamics that accounts for large oscillation amplitudes and nonplanar wakes is developed. An experiment with a cantilevered flat plate at low Reynolds number is set up and used to validate the low-order model, as well as to study post-flutter limit-cycle oscillations. Results from the low-order model are promising, but show that aerodynamic nonlinearities such as flow separation and leading-edge vortex shedding must also be modeled in order to predict all possible limit-cycle oscillations of the aeroelastic system.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All rights reserved.A simple and self-contained methodology to assess the validity of the assumption of small deformations in linear structural-dynamic models was recently proposed. The advantages of the methodology lie in the fact that it does not depend on the availability of higher-fidelity, nonlinear models: it is rather based on the selection of two different structural nodes where the structural motion is to be one at a time completely constrained, typically, a node near the center of mass and another in the region of maximum structural displacements with respect to mean axes. If the two displacement vectors calculated in each case can be transformed between themselves with linear rigid-body modes of the structure, then it is still in the regime of small deformations. In the present paper, in order to demonstrate the value of this methodology, it is applied to the X-HALE aircraft in its four-, six- and eight-meter-span configurations, and the results obtained with the assumption of small deformations are compared with a higher-fidelity model that comprises large structural deformations.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All Rights Reserved.Formulations for the flight dynamics of flexible aircraft have been commonly applied to aircraft free to fly in the three-dimensional space, having all six rigid-body degrees of freedom. For risk reduction in the future flight operations of the X-HALE testbed at ITA, however, wind-tunnel tests of the remotely-piloted, four-meter-span configuration of the aircraft were performed. In the wind tunnel, the rigid-body translations were completely constrained, but the same was not valid for the rigid-body rotations, which could be conveniently left free or not with a proper selection of the connection between the aircraft and the wind-tunnel mount. In the present paper, in order to computationally assess the response and stability characteristics of the aircraft in the wind tunnel, we derive equations of motion for a constrained flexible aircraft with up to three rigid-body rotational degrees of freedom, mounted on an also flexible wind-tunnel strut. The numerical model has its value confirmed by the wind-tunnel tests in the predicted and observed roll-control reversal for anti-symmetrical deflections of the all-moving tails, and absence of reversal for aileron deflections.

Copyright 2018, IADC/SPE Drilling Conference and Exhibition.An extension of Delayed Detached-Eddy Simulation (DDES) capabilities developed in SU2 to unsteady transonic buffet flow is present. An assessment of Spalart-Allmaras turbulence model variants with the 2D OAT15 airfoil reveals that the mixing layer compressibility correction plus the quadratic constitutive relation (SA-Comp-QCR) was the combination able to capture shock buffet accurately. Refined Roe scheme was also implemented, including adaptive dissipation function with Ducros shock sensor and Travin’s blending. The SU2 DDES implementation is tested in the Benchmark Supercritical Wing, analyzing the case 3 of the Second AIAA Aeroelastic Prediction Workshop. The results obtained are encouraging, showing good agreement for mean pressure coefficient and coherent fluid flow structures behind the shock.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Aircraft icing is a matter that still demands considerable research efforts because of its great impact on aircraft performance and safety. In most applications, the power source of Ice Protection Systems (IPS) is the engine, and thus fuel burn, engine thrust, and aircraft performance are affected by the compressed bleed air or shaft horsepower extracted. IPS pre-activation, intercycle, and residual ice shape and position in the leading edge are also important and do affect aircraft performance. This paper presents a proof of concept of a novel ice protection system based on Carbon Nanotubes (CNT) used as electrical heaters, installed in the leading edge of a two-dimensional horizontal tail model, and tested in an icing wind tunnel. The main advantages of the CNT heaters are their light weight, easiness to conform (very thin layer), and uniform electrical and thermal properties. The CNT based IPS model was tested in de-icing mode, except for a narrow zone along the leading edge highlight referred to as ‘parting strip’ that was operated in anti-icing mode. Based on the residual and intercycle ice accretions footprint obtained in the icing wind tunnel tests, the de-icing configuration tested was deemed successful. The associated aircraft performance degradation will be further investigated in on-going and future work by means of numerical analysis, wind-tunnel tests, flight tests with artificial ice shapes, and flight tests in natural icing conditions. In addition, future research will investigate the optimization of the de-icing system heating zones distribution (size and position) and de-icing sequence to potentially reduce the required power input or the residual and intercycle ice accretions.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.A detailed investigation of the shock wave formation patterns over a NACA0012 airfoil in a mostly laminar, transonic regime is presented. Information regarding the position of the boundary layer transition was obtained using Temperature Sensitive Paint (TSP) for similar conditions previously studied using Pressure Sensitive Paint (PSP). In this paper, previous results regarding shockwave formation patterns are revisited and further analyzed with the additional input of TSP data. Results indicate a strong correlation between the position of the boundary layer transition and the shockwave onset, confirming the hypothesis proposed in the previous investigation. Thus, the process of shockwave formation and the role of boundary layer interaction in it is further clarified. Besides Pressure and Temperature Sensitive Paints, the traditional method of pressure taps was used to confirm PSP measurements and account for possible deviations.

© Emerald Publishing Limited.Purpose: The purpose of this paper is to present a mathematical model of one very flexible transport category airplane whose structural dynamics was modeled with the strain-based formulation. This model can be used for the analysis of couplings between the flight dynamics and structural dynamics. Design/methodology/approach: The model was developed with the use of Hamiltonian mechanics and strain-based formulation. Nonlinear flight dynamics, nonlinear structural dynamics and inertial couplings are considered. Findings: The mathematical model allows the analysis of effects of high structural deformations on airplane flight dynamics. Research limitations/implications: The mathematical model has more than 60 degrees of freedom. The computational burden is too high, if compared to the traditional rigid body flight dynamics simulations. Practical implications: The mathematical model presented in this work allows a detailed analysis of the couplings between flight dynamics and structural dynamics in very flexible airplanes. The better comprehension of these couplings will contribute to the development of flexible airplanes. Originality/value: This work presents the application of nonlinear flight dynamics-nonlinear structural dynamics-strain-based formulation (NFNS-s) methodology to model the flight dynamics of one very flexible transport category airplane. This paper addresses also the way as the analysis of results obtained in nonlinear simulations can be made. Comparisons of the NFNS-s and nonlinear flight dynamics-linear structural dynamics methodologies are presented in this work.

© 2017.The stakeholder analysis is applied in numerous disciplines and may be called differently in each. In the discipline of Strategic Planning, the stakeholder analysis is performed when analyzing organizational environments. In the discipline of Project Management, a stakeholder analysis is performed for the feasibility of managing the Stakeholder Management Plan. In the discipline of Systems Engineering, the stakeholder analysis is performed during requirements analysis. In organizations that develop complex products, these three disciplines are relevant and activities of stakeholder analysis are repeated when each discipline is applied. This creates duplication and therefore waste. As a result, this article proposes to develop a unique method of stakeholder analysis that can achieve the goals of different stakeholders analyzes performed when applying the disciplines mentioned above. In general, it was concluded that the proposed method and its application show that a single stakeholder analysis can be performed to meet the objectives of Systems Engineering, Project Management and Strategic Planning, without the need for replication of analysis when applying each of these disciplines.

Copyright © 2016 by Flavio Silvestre. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.In this paper, the control law design for flexible aircraft is discussed. First, the traditional procedure of decoupling rigid-body and aeroelastic dynamics with low-pass and notch filters is addressed, with focus on controller performance as well as the resulting stability margin issues. A procedure based on a unified formulation of the flexible aircraft dynamics for flight control law design is proposed. In this procedure, the aeroservoelastic dynamics is assessed in the loop, and the offline filtering process is avoided. The formulation is applied to the virtual aircraft generic narrow-body airliner, with improvements in closed-loop performance and stability margins.

© 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.

© 2016 Elsevier LtdThis paper aims to estimate the modal parameters of composite flat plate models through Experimental Modal Analysis (EMA) using piezoelectric transducers. The flat plates are composed of three ply carbon-epoxy fibers oriented in the same direction. Five specimens with different unidirectional fiber nominal orientations θk (0o, 30o, 45o, 60o and 90o) were tested. These models were instrumented with one PZT (Lead Zirconate Titanate) actuator and one PVDF (Polyvinylidene Fluoride) sensor and an EMA was performed. The natural frequencies and damping factors estimated using only a single PVDF response were compared with the estimated results using twelve measurement points acquired by laser doppler vibrometry. For comparison purposes, the percentage error of each natural frequency estimation and the percentage error of the damping factor estimations were computed, as well as their averages. Even though the comparison was made between a SISO (Single-Input, Single-Output) and a SIMO (Single-Input, Multiple-Output) techniques, both results are very close. The vibration modes were estimated by means of laser measurements and were used in the modal validation. In order to verify the accuracy of the modal parameters, the Modal Assurance Criterion (MAC) was employed and a high correlation among mode shapes was observed.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.In spite of occurring during a short period of time, the transonic regime plays an important role in the flight envelope of an aerospace vehicle, given that complex flow phenomena take place at this speed range, as for example shock wave occurrences and associated shock waveboundary layer interactions (SBLIs). The present study is aimed at investigating the shock wave formation patterns over a NACA 0012 airfoil in the transonic regime as the Mach number and the angle of attack is varied. The angles of attack, α, of 0°, 2°, 4° and 5° were considered, and the Mach number, M, was varied between 0.2 and 0.8. In all tested conditions, measurements were conducted for very small variations of Mach number in the range close to shock wave occurrence, in order to study the shock wave patterns in this region and also to get insights on shockwave-boundary layer interactions (SBLIs) taking place. Significant differences both between laminar and turbulent shockwave formation patterns and fullydeveloped shock structures were detected, as well as Cpvariations due to laminar-turbulent transition.

Copyright © 2016 by ASME.In this paper we consider the flight dynamics of fighter aircraft at high angles of attack with uncertain aerodynamic coefficients. Stochastic parametric uncertainty is dealt with by employing spectral decomposition of the random variables by means of the generalized polynomial chaos expansion. We propose an optimal linear feedback strategy for the automatic pilot system to recover the aircraft from stall and provide acceptable dynamic response. Optimality of the proposed control law is proved by solving the Hamilton-Jacobi-Bellman equation and asymptotically stability of the controlled nonlinear aircraft model is guaranteed in the Lyapunov sense. Numerical results are verified with Monte-Carlo simulations.

Recent technological advances in mechatronics enhance the possibilities of utilizing subscale flight testing as a tool in the development of aircraft. This paper reports the current status of a joint Swedish-Brazilian research project aiming at exploring these possibilities. A 13% scale fighter aircraft is used as a test bench for developing methods and procedures for data acquisition. The aircraft is equipped with an instrumentation system assembled from off the shelf components as well as open source hardware and software.

© Copyright 2016 by the American Institute of Aeronautics and Astronautics, Inc.An inertially-coupled formulation for the flight dynamics of flexible aircraft undergoing small deformations is developed. The availability of a structural-dynamic finite element model of the aircraft is presupposed. With all the coupled dynamics taken into account, an arbitrary choice of the body reference frame can be made. This frame is also allowed to be noncoincident with the frame of reference used to calculate the aerodynamic loads. In the equations of motion, the inertial coupling terms are linearized in the elastic displacements around a calculated equilibrium condition. Appropriate modes of vibration are then used in the calculation of the dynamic deformation of the structure. A simple quasi-steady incremental aerodynamic model based on the vortex-lattice method is used. The formulation is tested in the flight simulation of an idealized forward-swept-wing aircraft model. Numerical results show that, under small deformations, different body axes lead to the same overall motion of the aircraft with respect to an inertial frame. The stiffness level at which geometrically nonlinear formulations would become necessary is also determined.

© 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.Nonlinear aeroelastic phenomena is continuously investigate in aeronautical researches. The nonlinearity nature, for these cases, could be aerodynamic, such as dynamic stall or shock waves or structural, for example, free play or large displacements due to high aspect ratio and high flexibility. This work will investigate a very flexible wing with high aspect ratio subjected to unsteady flow. A flutter analysis is proceeded in order to evaluate the error between the computational result and the experiment. Since the linear flutter theory consider small displacements, it is expected a nonlinear phenomena. In this case, the experiment time series is analyzed in order to understand this nonlinearity and a 0-1 test is performed to evaluate if the system has chaotic behavior.

This paper aims to improve the data acquisition in experimental aeroelastic analyses using PZT (Lead Zirconate Titanate) excitation and appropriate test parameters definition. High noise levels are usually associated to data acquisition in experimental aeroelastic analyses, leading to spurious roots in the stabilization diagram. Some procedures can be applied to improve the acquisition, for instance by working with the average number of a set of acquisitions. When working with the average number of a set of acquisitions, the noncorrelated inputs (such as noise) tend to disappear. Nevertheless, even when using a broader acquisition set from which to compute an average, it is not possible to obtain a good structural response, due to the poor relationship between signal and noise. Therefore, with the signal/noise relationship enhancement in mind an excitation can be used since it also helps to excite vibration modes that cannot be naturally excited by aerodynamic/turbulence forces. This study has four main investigation fields: (a) to choose the most appropriate type of signal excitation from the available ones; (b) to identify the average number, which improves the structural response when high speeds are the main focus of the wind tunnel test; (c) to define the test parameters aiming to improve the quality of the acquisition data, such as, estimator and window function; and (d) to identify the excitation level, which allows working in the linear structural response. Besides investigating the usage of PZT excitation to improve the data acquisition and showing the importance of an appropriate selection of the test parameters, this work also aims to encourage the development of investigative procedures to enhance data acquisition.

The present work deals with time response and stability simulations of flexible wings using a structural commercial solver, Abaqus. The use of commercial codes for structural analyses are necessary due to their ability to deal with more complex geometry than in-house codes usually do. The problem becomes establishing strategies for coupling these codes with the different aerodynamic solvers. In aeroelasticity this coupling involves passing displacements, velocities and forces from one model to another. Here, Abaqus is coupled to the commercial aerodynamic solver ZAERO. Initially, static stresses are computed based on loads obtained from trim analyses with ZAERO aiming establishing procedures for information coupling between codes. As a first approach, a simplified cantilever rectangular plate with a concentrated mass modelling a ballast located at the wing tip is simulated and then the results are correlated to experimental data for validation. Secondly, a more complex model of an aircraft is also analysed using the same methodology as described above. Finally, some conclusions are presented regarding the importance of including nonlinearity effects in the formulation of flexible wings and then compared to the same simulations assuming a linear theory. Final comments are also mentioned with guidelines for the computational procedures and limitations on the assumptions made to get reasonably accurate results.

Emergence of flutter compromises not only the long term durability of the wing structure, but also the operational safety, flight performance and energy efficiency of the aircraft. Effective means of flutter prevention are, therefore, mandatory in the certification of new flight vehicles. This work intends to address the application of viscoelastic material for flutter suppression in a typical section under quasi-steady and unsteady aerodynamic loads. A numerical procedure is proposed to solve the set of non-linear equations. A parametric study showing the influence of the temperature on the stiffness, damping and flutter velocity of the system is also carried out using different viscoelastic based damping arrangements. The preliminary results indicate that the aeroelastic behavior of the system is significantly affected by the temperature and viscoelastic damping arrangement.

The use of correction factors to improve the accuracy of the aerodynamic influence coefficient (AIC) matrices produced by the vortex-lattice and the doublet-lattice methods has been an engineering practice in the field of aeroelasticity. In order to account for either viscous or transonic flow effects not considered in the linearized formulation of such methods, the most frequent correction techniques have been to pre-multiply or to post-multiply the AIC matrices by diagonal matrices comprising semi-empirical weighting factors. This paper proposes a different correction approach: the control-point-placement method (CPPM), based on the idea of displacing the control point of each panel - the point where the boundary condition of flow tangency must be satisfied. Both the vortex- and the doublet-lattice methods have been developed with the singularities placed at the quarter-chord line of the panels and the control points at three quarters of their mean chords. With the calculation of modified control point positions, the CPPM intrinsically changes the mutual aerodynamic influence between the panels and allows the lifting surface methods to predict steady-state pressure distributions that match or approximate with minimum error those derived from wind tunnel measurements or higher-fidelity CFD solutions. Different approaches to extend the aerodynamic correction for application at non-zero reduced frequencies in the doublet-lattice method are then studied. Results are presented that are in acceptable agreement with benchmark wind tunnel data and comparisons are made between the proposed methodology and the traditional diagonal matrix corrections. © 2014 Published by Elsevier Masson SAS.

The use of tests is required when moving into areas sparsely explored by theory as an important tool for its validation. Aeroelastic wind tunnel tests using scaled models can be performed in order to verify the analytical methods, requiring a model that represents the problem qualitatively or, in a more complex case, checking the behavior of a real aircraft, requiring a representative model in which tests must necessarily be done demonstrating, statically and dynamically, their fidelity to the real structure. In this work, using previously acquired wind tunnel tests experimental data, a modal identification routine has been developed to analyse the data. Using theoretical scaled aircraft models, a theoretical versus experiment correlation was performed in order to verify the quality of the theoretical results.

This paper aims estimate modal parameters of thin composite wing models through experimental modal analysis (EMA) using piezoelectric materials. The wing models are flat plate based on three ply carbon-epoxy fiber in same directions. Five specimens with different unidirectional fiber nominal orientation (θk = 0°, θk = 30°, θk = 45°, θk = 60°, θk = 90°) are tested. These models were instrumented with one PZT (Lead Zirconate Titanate) actuator and one PVDF (Polyvinylidene Fluoride) sensor and results compared with vibrometer laser measurements. The orthotropic materials have different stiffness properties to each direction. These characteristics can be used to optimize or improve the dynamic behavior. Several examples, especially for aeronautic application can be cited, e.g., wings with negative deflection where fiber orientation was used to resolve divergence problem that these wings shows. An investigation about the dynamic behavior is conducted the modal parameters extracted. Using twelve points measured by vibrometer laser the vibration modes were achieved and modal assurance criterion (MAC) were calculated to assess the correlation between modes.

The current paper is concerned with studying the effects of using different unsteady computational fluid dynamics data in order to generate the root locus for aeroelastic stability analysis. The dynamic system being considered in the present work is a NACA 0012 airfoil-based typical section in the transonic regime. The CFD calculations are based on the Euler equations and the code uses a finite volume formulation for general unstructured grids. A centered spatial discretization with added artificial dissipation is used, and an explicit Runge-Kutta time marching method is employed. Unsteady calculations are performed for several types of excitation on the plunge and pitch degrees of freedom of the dynamic system. These inputs are based on step and orthogonal Walsh functions. The use of system identification techniques is employed to allow the splitting of the aerodynamic coefficient time histories into the contribution of each individual mode to the corresponding aerodynamic transfer function. Such transfer functions are, then, interpolated and used in an aeroelastic stability analysis in the frequency domain. The present work compares the results provided for each case and attempts to contribute with guidelines for such analyses. © 2012 AIAA.

The current paper is concerned with studying the effects of using different unsteady computational fluid dynamics data in order to generate the root locus for aeroelastic stability analysis. The dynamic system being considered in the present work is a NACA 0012 airfoil-based typical section in the transonic regime. The CFD calculations are based on the Euler equations and the code uses a finite volume formulation for general unstructured grids. A centered spatial discretization with added artificial dissipation is used, and an explicit Runge-Kutta time marching method is employed. Unsteady calculations are performed for several types of excitation on the plunge and pitch degrees of freedom of the dynamic system. These inputs are based on step and orthogonal Walsh functions. The use of system identification techniques is employed to allow the splitting of the aerodynamic coefficient time histories into the contribution of each individual mode to the corresponding aerodynamic transfer function. Such transfer functions are, then, interpolated and used in an aeroelastic stability analysis in the frequency domain. The present work compares the results provided for each case and attempts to contribute with guidelines for such analyses. © 2013 by J.H.A. Azevedo, J.L.F. Azevedo and R.G.A. Silva.

The paper is concerned with downwash correction methods for aeroelastic stability analyses in the transonic regime. The effects of the formulation used in the calculation of nonlinear, unsteady reference pressures are addressed, together with the influence of the motion amplitude. A finite-difference Euler/Navier-Stokes code is used to calculate the unsteady aerodynamic loading due to dynamic angle of attack variations in three-dimensional transonic flow. The computed unsteady pressure coefficients are used as a reference state for flutter analyses using the downwash weighting method. The test case considered is the well-known AGARD wing 445.6 standard aeroelastic configuration. The configuration is subjected to rigid body pitching oscillation about the mid-chord point at the root section. Flutter boundaries are computed using either inviscid or viscous-based unsteady pressures in the downwash correction methodology. The results are compared with available experimental data and they indicate that both viscous and thickness effects play an important role on the flutter prediction capability. © 2013 The Brazilian Society of Mechanical Sciences and Engineering.

Aeronautical engineering has faced a significant and continued development over the last decades towards the design of lighter, more maneuverable and more multidisciplinarily optimized aircraft, leading to more flexible vehicles. In this context, the fields of aeroelasticity and aeroservoelasticity play a very important and increasing role. Neglecting such flexibility effects on the flight dynamics and control system analysis and design may be an invalid premise, depending on how intense might be the coupling between the rigid and the flexible degrees of freedom. Traditional modeling approaches have often neglected the effects of inertial coupling in the treatment of the dynamics of the deformable aircraft, allowing great simplifications of the equations of motion. Most authors have indeed considered the body axes to be mean axes, what requires some care regarding the enforcement of the correct constraints and the expression of the aerodynamic force components along these axes directions. Looking for circumventing those limitations, while keeping the hypothesis of small local deformations, this work presents an integrated modeling methodology for the flight dynamics of deformable aircraft which takes into account all the coupled dynamics and is based on attached body axes. The formulation is developed for direct use with a finite-element model of the aircraft structure, with known distributed or lumped mass properties. The nonlinear inertial coupling terms are linearized with respect to the linear elastic displacements around an equilibrium condition. This condition is determined with the full nonlinear dynamics, considering displacement and load-transferal between the aerodynamic model and the finite-element model. Inertia-relieved constrained modes of vibration are then used as shape functions in the calculation of the dynamic deformation of the structure, thus not canceling the inertial coupling terms as would happen in the case of free-free normal modes. The proposed formulation is implemented and tested for simulating the flight of a generic narrow-body airliner (GNBA) model which has been developed for the purpose of these studies. The aerodynamic forces and moments are treated as the superposition of two contributions: the expected rigid-body ones and the incremental ones due to the structural deformation. The incremental aerodynamic forces and moments are modeled by the doubletlattice method (DLM). Rational-function approximation (RFA) together with the method of least squares for complex variables to determine the coefficients of the RFA and inverse Laplace transforms are employed to represent the reduced-frequency-domain forces in the time domain, leading to an augmented state-space system in which the aerodynamic lag phenomenon is taken into account.

The present work addresses a sensitivity analysis investigation of the aeroelastic stability margins for the VSB-30 sounding rocket during the atmospheric flight phase. Parametric stability analyses are performed considering variations of the inertia properties of the modular payload. Such variations can be caused by different type and/or number of experiments (payload modules). The aerodynamic model is based on a supersonic unsteady potential aerodynamic method. Freestream conditions depend on the flight speed and atmosphere. An equivalent structural dynamic model of the rocket is represented by a beam-like structure. The objective of this investigation is to establish an aeroelastic model for aeroelastic stability and response analyses, as well as a procedure for the identification of stability margins for rockets. The resulting aeroelastic model should be further used in MDO processes for the improvement of the vehicle flight performance. The results of the present effort indicate that the flutter behavior of the VSB-30 sounding rocket is sufficiently robust inside the operational envelope, even considering the environmental and loading conditions. The spinning effect, in this case, does not play a significant role, because the flutter margins remain almost unaltered with and without VSB-30 body spin.

The current paper is concerned with studying different forms of generating the aerodynamic operator, with the use of computational fluid dynamics (CFD) techniques, for performing transonic aeroelastic stability analyses in the frequency domain. The CFD calculations are based on the Euler equations and the code uses a finite volume formulation for general unstructured grids. A centered spatial discretization with added artificial dissipation is used, and an explicit Runge-Kutta time marching method is employed. The dynamic system being considered in the present work is a NACA 0012 airfoil-based typical section in the transonic regime. Unsteady calculations are performed for mode by mode and simultaneous excitation approaches. The simultaneous inputs used are based on orthogonal Walsh functions. The use of system identification techniques is employed to allow the splitting of the aerodynamic coefficient time histories into the contribution of each individual mode to the corresponding aerodynamic transfer functions. The present approach is validated against aerodynamic transfer functions obtained by indicial excitation of each individual mode. The results are in good agreement with the literature data and, hence, the procedure implemented accomplishes the desired goal of obtaining the aerodynamic operators for aeroelastic analyses with a single unsteady CFD calculation. ©2012 AIAA.

This paper presents a study on aeroelastic analyses of composite laminated wings subject to large displacements through the coupling of a nonlinear corotational shell finite element (FE) with an unsteady vortex-lattice method (UVLM) formulation. A FE implemented for the analysis of flat plates has been extended to model laminated composites with different lamina orientations. An UVLM formulation that is capable of coupling with this large displacement structural model is implemented. An explicit partitioned method is evaluated for the coupling of both models, using spline functions to interpolate information from the structural operator to the aerodynamic one, inside a Generalized-a time-marching solution. The resulting aeroelastic formulation provides a framework for the nonlinear aeroelastic analyses of structures made of composite material allowing the characterization of their nonlinear behavior and simulation of the limit-cycle oscillation response. Flat plate laminated wings designed for high flexibility and low flutter speed onset are used as investigation models. Effects of nonlinearities are easily observed in the numerical results, which are promising for expansion of the work and application to the analysis of more refined and complex composite flexible wings. ©2012 AIAA.

This work presents a structural optimization aided design methodology for composite laminated plates subject to fluid-structure interaction. The goal of the optimization procedure is to increase the flutter speed onset through the maximization of natural frequencies related to the vibration modes involved in the phenomenon. The aeroelastic stability analysis is performed using ZAERO software system, which includes ZONA 6 unsteady lifting surface method. The finite element method is applied to solve the structural model equilibrium equations, the eigenvalues sensitivities with respect to design variables are calculated analytically, and sequential linear programming is applied. The maximization is accomplished using two methods; the first method uses an aeroelastic analysis to determine which eigenmode causes the flutter onset, and its eigenvalue is then maximized. In the second method, a forward finite difference method is applied and the flutter speed sensitivities with respect to the eigenvalues are calculated. This sensitivity is used to guide the optimization process. Finally, a topology optimization problem is formulated to reduce the plate mass under a minimum flutter velocity constraint, using density distribution as the design variable. © Springer-Verlag 2012.

The current paper is concerned with studying different forms of generating the aerodynamic operator, with the use of computational fluid dynamics (CFD) techniques, for performing transonic aeroelastic stability analyses in the frequency domain. The CFD calculations are based on the Euler equations and the code uses a finite volume formulation for general unstructured grids. A centered spatial discretization with added artificial dissipation is used, and an explicit Runge-Kutta time marching method is employed. The dynamic system being considered in the present work is a NACA 0012 airfoil-based typical section in the transonic regime. Unsteady calculations are performed for mode by mode and simultaneous excitation approaches. The simultaneous inputs used are based on orthogonal Walsh functions. The use of system identification techniques is employed to allow the splitting of the aerodynamic coefficient time histories into the contribution of each individual mode to the corresponding aerodynamic transfer functions. The present approach is validated against aerodynamic transfer functions obtained by indicial excitation of each individual mode. The results are in good agreement with the literature data and, hence, the procedure implemented accomplishes the desired goal of obtaining the aerodynamic operators for aeroelastic analyses with a single unsteady CFD calculation. © 2012 by J.H.A. Azevedo, J.L.F. Azevedo and R.G.A. Silva.

Three simulation methodologies (LNS, NLAS and URANS) were compared with respect to capturing the flow field and the acoustic sources in an accurate and computationally cheap 2D simulation of the high lift airfoil, MDA 30P30N. Different grid methodologies were tested using the commercial code CFD++, by Metacomp. Initial tests analyzed the solution grid dependency on a series of 2D simulations and promising results for the URANS and NLAS simulations were obtained. For the 3D simulation, the commercial code PowerFLOW, by Exa, based on Lattice-Boltzmann equation, was used. This tool was available to the group relatively recently and, as yet, only a single mesh configuration was tested. Consistent vortical structures were captured on the slat cove and good agreement with PSD results from the literature was obtained for low frequencies. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

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 paper presents an expedient transonic correction technique to compute unsteady pressure distributions and aeroelastic stability in the transonic flow regime. The transonic correction procedure here is an improvement of the downwash weighting method proposed previously by several authors. The previous downwash weighting methods could provide pressure and/or force corrections to some extent by applying different weighting methods on the lifting-surface self-induced downwash resulting from aeroelastic structural displacements or prescribed motions. However, the resulting pressure/force solutions were often found to be inconsistent, because they all failed to include the proper transonic unsteady and out-of-phase effects. Our improved downwash correction method is a rational formulation to include proper transonic effects, as this formulation is based on a successive kernel expansion procedure established in accord with the formal pressure-downwash relation. Accordingly, the developed transonic correction procedure is a proper and rational one that is expected to yield more consistent aeroelastic solutions. This procedure is now a fully developed program, known as the transonic weighting aerodynamic influence coefficient procedure in the ZAERO software system, or ZTAW. Computed results by ZTAW for the unsteady pressures and aeroelastic stability boundaries for four selected wing planforms (AGARD 445.6, F-S, LANN, Lessing wings) are found to be in good agreement with measured data. In contrast to the computational-fluid- dynamics-based methods of computational aeroelasticity, the present procedure is proven to be far more computationally efficient and industrially viable while yielding comparable aeroelastic solutions. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The paper addresses further investigations on downwash correction methods for aeroelastic stability analyses in the transonic regime. The main concern is the investigation of the influence of the nature of the lifting surface motion, considering a general elastic body modal displacement to compute unsteady pressures. A finite-difference Navier-Stokes code is used to calculate the unsteady aerodynamic loads due to a three dimensional transonic flow. The unsteady pressure coefficients computed using this code are used as a reference state for flutter analyses based on a linearized aerodynamic theory using the downwash weighting method. The test case considered is the well-known AGARD wing 445.6 standard aeroelastic configuration. The results are compared with previous theoretical investigations.

The paper is concerned with downwash correction methods for aeroelastic stability analyses in the transonic regime. A finite-difference Navier-Stokes code is used to calculate the unsteady aerodynamic loading due to dynamic angle-of-attack variations in three-dimensional transonic flow. The computed unsteady pressure coefficients are used as a reference state for flutter analyses using the downwash weighting method. The effects of the amplitudes of motion used in the calculation of nonlinear, unsteady reference pressures are addressed. The test case considered is the well-known AGARD wing 445.6 standard aeroelastic configuration. The configuration is subjected to rigid-body pitching oscillation about the midchord point at the root section. Flutter boundaries are computed using unsteady pressures, in the downwash correction methodology, as reference conditions to compute weighting operators. The results are compared with available experimental data and they indicate that the aerodynamic interference and viscous and thickness effects play an important role on the flutter prediction capability.

A review of the current existing transonic AIC correction methods reveals that few can yield correct out-of phase pressures with a proper shock jump behavior. For this reason, a new method called the Transonic AIC Weighting (ZTAW) has been developed. This method consists of a downwash weighting matrix method derived from Giesing's force matching formulation in conjunction with a rational expansion of the classical kernel functions for lifting surfaces. Based on the general acceleration potential formulation in the frequency domain, a successive kernel expansion procedure has been established whereby the out-of-phase kernels can be successively expressed in terms of the in-phase kernels. In this way, the present method can yield correct out-of-phase pressures for elastic modes, with proper shock jump behavior, by only matching with given steady pressure inputs per one rigid pitch mode. Computed transonic cases include the F-5, LANN and Lessing wings for unsteady pressure validation and the AGARD445.6 and PAPA wings for flutter boundary validation. Good agreement with the measured unsteady pressures and the measured flutter boundaries assure the viability of the present method as an expedient industrial tool for transonic aeroelastic applications. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The paper is concerned with downwash correction methods for aeroelastic stability analyses in the transonic regime. The effects of the formulation used in the calculation of nonlinear, unsteady reference pressures are addressed, together with the influence of the motion amplitude. A finite-difference Euler/Navier-Stokes code is used to calculate the unsteady aerodynamic loading due to dynamic angle of attack variations in three dimensional transonic flow. The computed unsteady pressure coefficients are used as a reference state for flutter analyses using the downwash weighting method. The test case considered is the well-known AGARD wing 445.6 standard aeroelastic configuration. The configuration is subjected to rigid body pitching oscillation about the mid-chord point at the root section. Flutter boundaries are computed using either inviscid or viscous-based unsteady pressures in the downwash correction methodology. The results are compared with available experimental data and they indicate that both viscous and thickness effects play an important role on the flutter prediction capability.

The aeroelastic behavior of an aircraft is typically more critical in the transonic flight regime. The linearized potential based equations of the fluid flow do not allow accurate predictions of transonic flutter. Downwash weighting methods are adequate tools for approximating the nonlinear behavior of unsteady transonic flows, as far as aeroelastic applications are concerned. The purpose of downwash weighting methods is to correct unsteady pressures computed from linear aerodynamic models, to take into account nonlinear effects. Such methodology is less expensive than time domain computational aeroelasticity simulations. The reduced frequency to be considered in those methods is associated to specific conditions to be investigated for aeroelastic stability. This implies in a less expensive computational simulation, since it may be performed on a single harmonic motion simulation at the specific reduced frequency. The objective of the present work is to perform a sensitivity study with regard to the variation of the dynamic amplitude of the prescribed mode shape. A set of amplitudes of the disturbance in angle of attack is used to generate the non-linear unsteady pressure data. Therefore, it is possible to understand the variation of the computed flutter speeds with respect to the nature of the non-linear unsteady pressures. The flutter speed computation presents significant variations with respect to the nature of the unsteady pressure data. The results presented herein regarding the dependence of the flutter speeds on the amplitude of the motion indicate that the downwash correction method is closely related to the magnitude of such displacements.

Navier-Stokes code was used to verify the locally linear behavior of aerodynamic loads with respect to the dynamic angle of attack. Viscous and three-dimensional effects were taken into account and found to be significant.

A finite-difference Navier-Stokes code is used in order to study the linearity of aerodynamic loads with respect to the dynamic angle of attack in threedimensional transonic flow. Steady and unsteady pressure coefficients for prescribed rigid angle of attack motion are computed for a F-5 wing for which the method has been previously validated. The study is aimed at identifying the conditions under which approximate flutter analyses based on corrections to aerodynamic influence coefficients may be used. Results indicate that higher harmonics of the unsteady loads are present in the aerodynamic response and that the boundaries for linear behavior depend on the spanwise location along the wing. © 2002 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc.