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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.

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.

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.

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. 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. 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. 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.

The procedures to be performed after sudden engine failure of a single-engine helicopter impose high workload on pilots. The maneuver to regain aircraft control and safe landing is called autorotation. The safety limits to conduct this maneuver are based on the aircraft height versus speed diagram, which is also known as "Dead Man’s Curve”. Flight-test pilots often use subjective methods to assess the difficulty to conduct maneuvers in the vicinity of this curve. We carried out an extensive flight test campaign to verify the feasibility of establishing quantitative physiological parameters to better assess the workload endured by pilots undergoing those piloting conditions. Eleven pilots were fully instrumented with sensors and had their physiological reactions collected during autorotation maneuvers. Our analyses suggested that physiological measurements (heart rate and electrodermal activity) can be successfully recorded and useful to capture the most effort-demanding effects during the maneuvers. Additionally, the helicopter’s flight controls displacements were also recorded, as well as the pilots’ subjective responses evaluated by the Handling Qualities Rate scale. Our results revealed that the degree of cognitive workload was associated with the helicopter’s flight profile concerning the Height-Speed diagram and that the strain intensity showed a correlation with measurable physiological responses. Recording flight controls displacement and quantifying the pilot's subjective responses show themselves as natural effective candidates to evaluate the intensity of cognitive workload in such maneuvers.

Among 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.

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.

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., Microstates are feasible neural correlates in challenging and intrinsically noisy environments. Microstate templates recovered in a challenging experimental condition resemble previous reports in controlled environments. Microstate dynamics is modulated by cognitive activities. In particular, piloting activity during the maneuver to recover a falling helicopter does not modulate Microstate D, traditionally associated with the cognitive control and attention networks, but Microstate F, which involves the medial prefrontal cortex. This microstate class is potentially associated with the emotional regulation network.

Small unmanned aerial vehicles (UAVs) face flying quality problems different from those encountered by larger aircraft. The lower airspeeds and small dimensions make these vehicles more susceptible to gusts and stability and control issues, which may render the aircraft difficult to fly. Moreover, due to many factors, UAVs are often built with a considerable degree of uncertainty regarding their aerodynamic properties and flying quality. The resulting aircraft may present poor stability and handling characteristics. This work presents the conceptual design of a robust stability augmentation system (SAS), aimed at increasing stability characteristics and protecting aircraft prone to flying quality problems. In order to deal with parametric uncertainties, the controller was designed with the robust H-infinity technique. The design process is presented, and a parametric aircraft model is provided, together with longitudinal stability and control derivatives. Simulations are presented to show the effects of the controller on the aircraft behavior.

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.

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.

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.

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 wit...

Temperature-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.

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.

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.

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.