Recently, Wilson et al. [1] suggested the use of an active hinged folding wingtip device on aircraft wings, with the goal of benefiting from the aerodynamic efficiency of high-aspect ratio wings while reducing the peak loads that are experienced at the wing root in the presence of gusts. The dynamic load reduction potential of this approach has been demonstrated successfully in several studies, e.g., in [2,3]. In these studies, the timing of the hinge release with respect to the gust encounter has been identified as an important performance parameter, motivating further research on this topic. So far, the experimental data from wind tunnel measurements on the hinged folding wingtip that are available in the published literature are limited to only a few parameters, such as the wing root bending moment or the wingtip fold angle. In this study, the aeroelastic gust response of a hinged folding wingtip device will be analyzed based on a wind tunnel experiment performed at Delft University of Technology (see Fig. 1). The measurements in the wind tunnel are conducted with an integrated optical approach [4] that provides measurements of the dynamic response of the structure via tracking optical markers, and for the first time, of the flow field around the folding wingtip. The goal of this research is to provide detailed insights into the aeroelastic gust response of the folding wingtip at different hinge release timings and to produce reference data for numerical prediction models, in particular with respect to the distribution of the unsteady lift force acting on the folding wingtip.@en

The effects of the wing skin distortion on the boundary layer of a highly flexible wing are analyzed in a wind tunnel experiment using infrared thermography measurements. Considerable differences in the boundary layer flow are observed when comparing the sections of the wing near the ribs, where the design shape of the wing is preserved, and in between the ribs. At the spanwise locations between the ribs, the sectional wing shape distorts and triggers boundary layer transition close to the leading edge. The differences between the design behavior of the wing and the experimental results of the boundary layer analysis demonstrate the need for considering the skin deformation and its effects on the boundary layer flow when designing highly flexible wings.

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The laminar separation bubble (LSB) that forms on the suction side of a modified NACA 64 ₃-618 airfoil at a chord-based Reynolds number of Re = 200 , 000 is studied using wind tunnel experiments. First, the LSB is characterized over a range of static angles of attack, in terms of the locations of separation, transition and reattachment—using surface pressure measurements, particle image velocimetry (PIV) and infrared thermography (IT). For the conditions tested, excellent agreement between the techniques is obtained. Subsequently, a pitching motion is imposed on the wind tunnel model, with reduced frequencies up to k = 0.25. While surface pressure measurements and PIV are not affected by the change in experimental conditions, the infrared approach is impaired by the thermal response of the surface. To overcome this, an extension of the differential infrared thermography (DIT) method for detecting the three characteristics of an unsteady LSB is considered. All three experimental techniques indicate a hysteresis in bubble location between the pitch up and pitch down phases of the motion, caused by the effect of the aerodynamic unsteadiness on the adverse pressure gradient. However, the DIT measurements suggest a larger hysteresis, which is attributed to the thermal response time of the model surface. The experimental results measured with the pressure sensors reveal that the hysteresis in bubble location is larger than the hysteresis in lift, indicating that the observed bubble hysteresis is not purely due to instantaneous flow conditions, but has an inherent component as well.

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The Pazy Wing test case is a benchmark for the investigation of aeroelastic effects at very large deflections. Tip deformations in the order of 50% span were measured in wind tunnel tests which renders this model highly attractive for the validation of numerical aeroelastic methods and tools for geometrically nonlinear, large deflection analyses. The present work is focused on high fidelity aerodynamic and aeroelastic simulations of the wing using RAN and URANS with transition modeling in order to capture nonlinear effects originating from the shape of the wing and the low Reynolds number. Steady and unsteady aerodynamic as well as static coupling simulations with a nonlinear structural model are presented, the impact of the different transition and turbulence modeling techniques is depicted. This work supports the Large Deflection Working Group (LDWG), which is one of the sub-groups of the 3rd Aeroelastic Prediction Workshop (AePW3). The work of NASA to generate a highly accurate geometry file for the generation of the CFD grid based on the scanned outer surface of the TU Delft Pazy Wing is gratefully acknowledged. Furthermore, the help of the members of the Large Deflection Working Group of the AePW3 is appreciated, especially for the setup of the beam model (Cristina Riso and Bret Stanford) as well as the experimental data (Arik Drachinsky and Daniella Raveh). Also the support of the DLR Institute of Aeroelasticity for this work is acknowledged.

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This study presents an aeroelastic wind-tunnel experiment to identify the influence of the wing stiffness and hinge release threshold on the gust load alleviation performance of a folding wingtip design. Five models with different stiffness and tailoring properties are tested, and the wing root bending moment at different conditions is compared to the response with the locked-hinge condition to assess the impact on the gust load alleviation capabilities of the folding wingtip. The results show that the structural properties do not have an important impact on the peak load alleviation but the hinge release threshold and timing do. Releasing with the correct timing can reduce significantly the peak loads. However, the dynamics of the system are affected by this release; the flutter speed is decreased, and, although the performance can improve, load oscillations increase, which can be considered detrimental for reasons such as fatigue or passenger comfort.@en

This thesis presents a novel measurement approach for aeroelastic wind tunnel testing. The key novelty of this approach is the integrated measurement of aerodynamic and structural quantities using an optical technique. The considered approach consists of combined measurements of flow tracer particles and structural markers using a Lagrangian particle tracking system. Based on these measurements, the quantities of interest for the characterization of an aeroelastic interaction, which are the three forces in Collar’s triangle (aerodynamic, elastic, and inertial), are determined. Currently, measurements in aeroelastic wind tunnel tests are typically performed with individual sensors for each quantity of interest (pressure transducers, strain gauges, or accelerometers) that are installed inside the experimental model and/or with a force balance that measures the total loads acting on the model. The integrated optical measurement approach is an advancement over this existing measurement technology because it provides field measurements of the aeroelastic structural response and the unsteady flow field around the experimental model, based on which the aerodynamic and structural load distributions can be determined, without requiring an instrumentation of the model with sensors. This measurement approach is therefore an effective way to produce experimental reference data to support the development of novel aeroelastic prediction methods with a potential to accelerate the technological development process for innovations in aeronautics in the future. The development and applications of the integrated optical measurement in this thesis are based on the measurements that were performed in three experimental campaigns in the wind tunnel. Each of the three experiments corresponds to one of the three main chapters of this thesis. All three experiments are performed on a largemodel scale, with dimensions on the order of 1m, which is a scale of high practical relevance for aeroelastic wind tunnel testing. The complexity of the three experiments, in terms of the aeroelastic phenomena that are observed, is increased incrementally, from a rigid-body motion, over a linear aeroelastic test case, to a nonlinear aeroelastic test case. Based on the observations and findings of the previous experiments, the data analysis methods for the subsequent experiments are selected and applied. The first measurements with the integrated approach...@en

The aeroelastic response of the Delft-Pazy wing to steady and periodic unsteady inflow conditions is analyzed experimentally. The Delft-Pazy wing is a highly flexible wing model based on the benchmark Pazy wing (Avin, O., Raveh, D. E., Drachinsky, A., Ben-Shmuel, Y., and Tur, M., “Experimental Aeroelastic Benchmark of a Very Flexible Wing,” AIAA Journal, Vol. 60, No. 3, 2022, pp. 1745-1768) and exhibits wingtip displacements of more than 24% of the span in the present study. The nonintrusive measurements are performed with an integrated optical approach that provides combined measurements of the structural response of the wing and the unsteady flowfield around it. The aeroelastic loads acting on the wing are derived using physical models and validated against force balance measurements, showing a good agreement for all considered inflow conditions. The analysis of the aeroelastic response of the wing to the unsteady inflow produced by a gust generator shows that both structural and aerodynamic responses depend strongly on the frequency of the gust. The results of this study provide a characterization of the aeroelastic behavior of the Delft-Pazy wing and can serve as a reference for the development of novel and improved nonlinear aeroelastic simulation models.

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An aeroelastic wind tunnel experiment to identify the influence of the wing stiffness and hinge release threshold on the gust load alleviation performance of a folding wingtip design is presented in this study. Five models with different stiffness and tailoring properties are tested and the wing root bending moment at different conditions is compared to the response with locked hinge conditions to assess the impact on the gust load alleviation capabilities of the folding wingtip. The results show that the structural properties do not have an important impact on the peak load alleviation but the hinge release threshold and timing do. Releasing with the correct timing can reduce significantly the peak loads. However, the dynamics of the system are affected by this release: the flutter speed is decreased and, although the performance can improve, load oscillations increase, which can be considered detrimental for reasons such as fatigue.@en

The aeroelastic response of a highly flexible wing to periodic gust excitation is determined experimentally. The integrated optical measurement approach that is applied provides combined measurements of the structural response of the wing and the unsteady flow field around it. The aeroelastic loads acting on the wing are derived from these measurements using physical models and validated against force balance measurements. It is observed that both structural and aerodynamic responses to a periodic gust excitation of a given amplitude depend strongly on the frequency of the gust. The obtained data set of results provides a complete description of the aeroelastic response that is suited as a reference for the development of aeroelastic simulation models.@en

The aerodynamic, elastic, and inertial components in Collar’s triangle of forces acting on a flexible wing with span width s=1.75  m and chord c=0.25  m are determined based on integrated optical measurements of the structural and aerodynamic response to steady and unsteady periodic inflow conditions at a chord-based Reynolds number of 2.3×105. The measurement device is a coaxial volumetric velocimeter mounted on a robotic arm, which is used to perform optical measurements of fiducial markers on the wing surface, and helium-filled soap bubbles, which are used as flow tracers. The optical measurements of the structural markers and the flow tracers are both processed with the Lagrangian particle tracking algorithm Shake-the-Box. Subsequently, physical models are used to determine the inertial and elastic forces of the aeroelastic interaction from the marker tracking results, and to determine the unsteady aerodynamic lift force from the flow velocity fields. The results of this integrated aeroelastic characterization approach are in physical agreement with each other according to the equilibrium of forces in Collar’s triangle and good agreement with external reference measurements.@en

The aeroelastic response of a very flexible wing in steady and unsteady inflow conditions is measured in a wind tunnel experiment. An integrated aeroelastic characterization is performed with a non-intrusive optical setup that allows simultaneous measurements of the structural motion and the flow field around the wing. The experimental aerodynamic loads that are inferred from the flow field measurements are in very good agreement with reference data from a force balance. Prior to the wind tunnel experiment, results of the numerical modal analysis of a structural model of the wing are compared with the experimental results from a ground vibration test. An aeroelastic model validation is achieved by applying the experimental aerodynamic loads to the structural model of the wing. The results of this structural analysis are compared with the measured structural response in the wind tunnel for steady inflow conditions, yielding differences of around 15% in tip displacements when using a linear structural model.@en

The Pazy Wing test case is a benchmark for the investigation of aeroelastic effects at very large deflections. Tip deformations in the order of 50% span were measured in wind tunnel tests, which renders this model highly attractive for the validation of numerical aeroelastic methods and tools for geometrically nonlinear, large deflection analyses. Due to the low flow velocity (up to 60 m/s) and the simple geometry of the wing, simulation programs based on subsonic, linear potential aerodynamic solvers (such as VLM and UVLM) are an ideal basis for static coupling and flutter simulations. However, more comprehensive analyses with focus on aerodynamic nonlinearities such as stall and limit cycle oscillations (which have been observed in several wind tunnel tests) are attractive research topics but call for advanced aerodynamic methods. The present work is thus focused on high fidelity aerodynamic simulations of the Pazy Wing using RANS with transition modeling in order to capture nonlinear effects originating from the particular shape of the wing and the low Reynolds number. It is a collaborative activity of DLR, NASA, and TU Delft and supports the Large Deflection Working Group (LDWG), which is one of the sub-groups of the 3rd Aeroelastic Prediction Workshop (AePW3).@en

The unsteady surface pressure distribution and aerodynamic loads on a pitching airfoil are determined non-intrusively using PIV measurements. An experimental test case is considered where the flow around the airfoil is mostly attached while the unsteady effects on the aerodynamic loads are significant. The surface pressure is calculated from the flow velocity measurements in the vicinity of the airfoil surface, that are obtained with a robotic PIV system, by using relations from unsteady potential flow and thin airfoil theory. The proposed approach is a robust and computationally efficient approach to obtain non-intrusive measurements of the unsteady surface pressure distribution and the aerodynamic loads, that are in good agreement with reference data from installed pressure transducer sensors.

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The three components in Collar’s triangle of forces (aerodynamic, elastic and inertial) acting on a flexible wing are determined, based on integrated optical measurements of the structural and the aerodynamic response to steady and unsteady periodic inflow conditions. The measurement device is a coaxial volumetric velocimeter mounted on a robotic arm, which is used to perform optical measurements of fiducial markers on the wing surface and simultaneously of helium-filled soap bubbles that are used as flow tracers. The optical measurements of the structural markers and the flow tracers are both processed with the Lagrangian particle tracking algorithm Shake-The-Box. Subsequently, physical models are used to determine the inertial and the elastic force of the aeroelastic interaction from the marker tracking results, and to determine the unsteady aerodynamic lift force from the flow velocity fields. The results of this integrated aeroelastic characterization approach are in physical agreement with each other according to the equilibrium of forces in Collar’s triangle and good agreement with external reference measurements.@en

The structural motion and unsteady aerodynamic loads of a pitching airfoil model that features an actuated trailing edge flap are determined experimentally using a single measurement and data processing system. This integrated approach provides an alternative to the coordinated use of multiple measurement systems for simultaneous position and flow field measurements in large-scale fluid–structure interaction experiments. The measurements in this study are performed with a robotic PIV system using Lagrangian particle tracking. Flow field measurements are obtained by seeding the flow with helium-filled soap bubbles, while the structural measurements are performed by tracking fiducial markers on the model surface. The unsteady position and flap deflection of the airfoil model are determined from the marker tracking data by fitting a rigid body model, that accounts for the motion degrees of freedom of the airfoil model, to the measurements. For the determination of the unsteady aerodynamic loads (lift and pitching moment) from the flow field measurements, two different approaches are evaluated, that are both based on unsteady potential flow and thin airfoil theory. These methods facilitate an efficient non-intrusive load determination on unsteady airfoils and produce results that are in good agreement with reference measurements from pressure transducers.

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The aerodynamic loads on a flexible wing in terms of the surface pressure distribution and the lift force along the span are determined experimentally based on non-intrusive Lagrangian particle tracking (LPT) measurements. As the flexible wing deforms under the aerodynamic loads, its deformed shape is first reconstructed based on structural LPT measurements conducted together with the flow measurements in an integrated approach. Based on the reconstructed wing shape, flow tracers data are collected along surface normals to evaluate the surface pressure, as well as along elliptic paths around the wing to determine the circulation. The lift force is calculated from the surface pressure by integrating the pressure difference along the chord, as well as from the circulation using the Kutta-Joukowski theorem. The circulation-based lift results are in very good agreement with reference measurements from a force balance, with differences in the total lift force on the wing of less than 5%. The lift estimation based on the extrapolated surface pressure is consistently lower than the circulation-based lift, by about 10%, due to the limited accuracy of the pressure extrapolation near the leading edge region, where a considerable fraction of the lift is generated.@en

A new approach to measuring unsteady boundary layer transition in periodic processes with infrared thermography is introduced. The radiation from the heated suction surface of a pitching airfoil model is measured with an infrared camera. The extraction of the extrema of the measured radiation signal at fixed model locations yields instants of the motion phase that correlate with the occurrence of boundary layer transition. The analysis of the extrema of the signal’s gradient produces results that are equivalent to the results acquired by optimized differential infrared thermography, and it provides insight into the origins of the systematic error of that technique. It is demonstrated that the local infrared thermography approach can be readily extended to measuring two-dimensional boundary layer transition fronts.@en

A proportionate controller is investigated experimentally for unsteady load alleviation purposes on a 2D wing model with a trailing-edge flap. The controller acts on the velocity of the flaps, and pressure sensors are used to detect the unsteady loads, which are generated by actuating the wing model in a sinusoidal motion. Two different regimes are considered: attached flow and dynamic stall. The influence of actuation frequency and controller time lag is also studied. A reduction of 87.5% in the standard deviation of the lift is obtained for a frequency of 0.2Hz and time lag in the control system of 12ms for attached flow conditions. The reduction of the standard deviation of the lift deteriorates for increased frequency and time lag. The proposed controller is also able to reduce the loads during dynamic stall, although the reduction is smaller, close to 40%, and can negatively affect the aerodynamic damping of the model. The flap actuation is also shown to delay the onset of dynamic stall, by increasing the static stall angle with respect to the case without flap deflection.

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Advanced data processing methods for detecting unsteady boundary layer transition in periodic aerodynamic processes by means of infrared thermography measurements are presented. The thermal radiation emitted from the heated suction surface of a pitching airfoil model in subsonic flow is measured with an infrared camera. The unsteady boundary layer transition location is detected by analyzing the difference in the infrared radiation signal over short periods of time with differential infrared thermography (DIT). The DIT method is optimized and automated in the present study, which facilitates the extension of the part of the motion period where valid DIT transition measurements are produced. Additionally, a new infrared thermography data processing method is introduced in this study. The extraction of the extrema of the measured radiation signal at fixed locations on the model surface yields instants of the motion period that relate to the occurrence of boundary layer transition. The local infrared thermography (LIT) approach can be extended to measuring the two-dimensional unsteady boundary layer transition front.@en

Abstract: Differential infrared thermography (DIT) is a method of analyzing infrared images to measure the unsteady motion of the laminar–turbulent transition of a boundary layer. It uses the subtraction of two infrared images taken with a short-time delay. DIT is a new technique which already demonstrated its validity in applications related to the unsteady aerodynamics of helicopter rotors in forward flight. The current study investigates a pitch-oscillating airfoil and proposes several optimizations of the original concept. These include the extension of DIT to steady test cases, a temperature compensation for long-term measurements, and a discussion of the proper infrared image separation distance. The current results also provide a deeper insight into the working principles of the technique. The results compare well to reference data acquired by unsteady pressure transducers, but at least for the current setup DIT results in an additional measurement-related lag for relevant pitching frequencies. Graphical abstract: [Figure not available: see fulltext.]

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