Structural Analysis and Experimental Tests of a Morphing-Flap Scaled Model

Abstract

The implementation of morphing wing mechanisms shows significant potential for improving aircraft performance, as highlighted in the recent literature. The Clean Sky 2 AirGreen 2 European project team is currently performing ground and wind tunnel tests to validate improvements in morphing wing structures. The project aims to demonstrate the effectiveness of these morphing designs on a full-scale flying prototype. This article describes the design methodology and structural testing of a scaled morphing-flap structure, which can adapt to three different morphing modes for various flight conditions: low-speed (take-off and landing) and high-speed (cruise). A scale factor of 1:3 was selected for the wind tunnel test campaign. Due to challenges in scaling the embedded mechanisms and actuators necessary for shape-changing, a full geometrical scale of the real flap prototype was not feasible. Static analyses were performed using the finite element method to address critical load conditions determined through three-dimensional computational fluid dynamic (CFD) analysis. The finite element (FE) analysis was conducted and the results were compared with the empirical data from the structural test. Good correlations were found between the structural testing results and numerical predictions, including static deflections and elastic deformations under applied loads. This indicates that the modeling approaches used during the design and testing phases were highly successful. Based on simulations for the ultimate load conditions tested during the wind tunnel tests, the scaled flap prototype has been deemed suitable for further testing.

1. Introduction

The concept of adjusting aircrafts’ geometry to optimize aircraft performance across various flight conditions has long been a key goal in aviation. Traditionally, this has been achieved through movable components such as ailerons, slats, and flaps, which help extend an aircraft’s operational capabilities [1]. Despite their benefits, these conventional solutions have drawbacks, including added weight from actuators and disruptions to aerodynamic surfaces. To address these challenges, researchers have explored alternative design approaches, with morphing technology being a prominent example [2]. Morphing designs focus on developing actuation techniques that enable smooth and significant changes in an aircraft’s shape. This approach offers the advantage of altering curvature substantially without the complexities associated with traditional slats and flaps, such as issues with gap overlap and deflections. Recent studies have extensively investigated various morphing technologies to identify the most efficient solutions for integrating these innovations into aircraft design [3].

In the aviation industry, high-lift systems on aircraft wings have been widely adopted for managing lift and drag forces during critical phases like take-off and landing. Traditionally, these systems use discrete and rigid components like leading-edge slats and trailing-edge flaps, which are mechanically adjusted via hinges and linkages to alter wing geometry. However, this conventional method amplifies system complexity and overall structural weight. Additionally, the presence of flap slots and their accompanying fairings has been identified as a major source of both airframe noise and aerodynamic drag [4].

Throughout aviation history, particularly from the 1910s and 1920s, there have been several instances of manned aircraft design innovations. One early example [5] involves a patent for kinematic systems that altered the external configuration of the wing leading and trailing edges. Another significant development allowed for the self-adaptation biplane and triplane profiles [6]. A major advancement came with the F-111, which featured wings capable of continuously adjusting their leading and trailing-edge cambers. The extensive testing of the F-111 demonstrated its enhanced capabilities in loitering, diving, and maneuvering, showcasing the potential of adaptive aerodynamic technologies [7,8]. Additional research at the Research Laboratory in Active Controls, Avionics, and Aeroservoelasticity Laboratory (LARCASE) at École de technology supérieure (ÉTS) has advanced morphing winglet and horizontal tail systems, specifically for the CRJ-700 [9,10,11] and the Cessna Citation X business aircraft [12,13].

In recent decades, morphing technologies have seen substantial and promising advancements, including innovative morphing mechanisms, novel actuation methods, and the use of smart materials. One notable development is the belt-rib concept, which was developed and tested at German Aerospace Centre (DLR) as part of the Adaptive Wing project. This concept features a structural framework, consisting of a closed shell, known as the belt, reinforced by in-plane stiffening elements called spokes. These spokes are typically connected to the belt via solid-state hinges [14]. Experimental studies have shown that the belt-rib concept is a viable and promising solution for controlling airfoil shape and achieving geometric adaptability [15].

Another morphing trailing-edge (TE) concept was developed and tested under the DARPA/AFRL/NASA Smart Wing program. This initiative aimed to demonstrate the feasibility of using smart-material-based actuators to create control surfaces for adaptive wing structures that eliminate gaps and hinges. The project’s goal was to improve aerodynamic and aeroelastic performance of military aircraft. The flexible skin–flexcore concept, a key component of this program, featured three main elements: an outer skin of elastomeric silicone, a flexible honeycomb structure, and a central leaf spring at its core [16].

Wind tunnel experiments have demonstrated that smart trailing-edge control surfaces can be tilted up to 20 degrees in less than 0.33 s. These movements were tested across various spanwise shapes, including uniform and nonuniform configurations such as linear and quadratic ramps, sine waves, and cosine waves. Daynes and Weaver [17] advanced this research with a new flap design. Their design featured a combination of components: an upper skin made of compliant carbon-fiber-reinforced plastic (CFRP), a flexible honeycomb core, and a lower skin composed of flexible silicone material. Their analysis revealed that this morphing flap could reduce tip deflection by 30% while still producing a lift change equivalent to that of a conventionally hinged flap with the same chord length [18].

Another innovative camber morphing concept, known as Fish Bone Active Camber (FishBAC), was introduced by Woods and Friswell [19]. This design allows for significant adjustments in camber in both directions and shows potential for various applications, including fixed-wing aircraft, helicopters, tilt-rotor aircraft, and wind turbines.

The design integrates four key components that together facilitate a robust and adaptable camber morphing mechanism. These components are a flexible backbone with stringers, a pre-tensioned elastomeric matrix composite (EMC) skin, an opposing tendon drive system with a locking spooling pulley to prevent back driving, and a rigid, non-morphing main spar.

Hetrick et al. conducted a study on FlexSys Inc. (Arbor, MI, USA)’s Mission Adaptive Compliant Wing (MACW), which focuses on a morphing trailing-edge concept integrated and flight-tested on a large manned aircraft. The study demonstrated that the morphing flap significantly enhanced aerodynamic performance during flight tests [20]. The MACW effectively reduced separated flow and drag across various lift conditions by enabling smooth trailing-edge deflections via an internal servo mechanism. Flight tests conducted at high speeds and altitudes highlighted the MACW’s potential for endurance flight applications, achieving a morphing rate of 30° per second at Mach 0.4. The study projected fuel savings of up to 15% for aircraft equipped with MACW-like wings, emphasizing the reduced actuation force and power requirements compared to conventional flaps. The research suggests that morphing wing concepts, like the MACW, could offer promising alternatives to traditional control surfaces [21].

FlexSys Inc. further advanced its technology by developing FlexFoil™, a transformative control surface capable of achieving notable deflections between −9° and +40° by leveraging distributed compliance enabled by a unique internal mechanism. The company posits that incorporating FlexFoil™ in extended-range aircraft could reduce drag by 5% to 12%, significantly improving fuel efficiency. Furthermore, FlexFoil™ ensures a smooth and continuous integration with the rest of the wing structure, even during deflection [22]. The FlexSys technology was implemented and flight-tested on NASA’s Gulfstream-III Subsonic Research Aircraft Testbed (SCRAT) as part of the Adaptive Compliant Trailing Edge (ACTE) project. In this evaluation, the SCRAT aircraft was equipped with the FlexSys morphing trailing edge, replacing the conventional aircraft flaps. The primary goal of this evaluation was to assess the aerodynamic efficiency and noise reduction capabilities of the compliant morphing trailing edge [23].

The Flexsys Mission Adaptive Compliant Wing represents a significant advancement in aircraft performance and efficiency. However, integrating MACW systems into commercial aviation necessitates compliance with certification standards set by regulatory authorities such as the Federal Aviation Administration (FAA) and the European Union Aviation Safety Agency (EASA). The certification process, however, presents challenges related to structural integrity, material fatigue, safety standards, and overall regulatory compliance. The utilization of advanced materials and flexible structures demands a re-evaluation of traditional testing methodologies to ensure the longevity and fatigue resistance of these novel materials across various operational scenarios.

The distinction between military and civil aviation is essential due to the notable differences in challenges and requirements in each domain. Consequently, the solutions and obstacles to be tackled vary significantly. In the civil sector, additional differentiations emerge based on the scale of the aircraft, ranging from large jets to commuter planes and general aviation. These classifications can be further divided into specific types, such as tourism, acrobatic, and ultralight aviation, each with distinct characteristics and considerations.

Despite significant efforts, the scientific and technological community is beginning to question why morphing technology has not yet matured enough for practical application, particularly in larger aircraft. Currently, the European community has invested significant resources in various research initiatives aimed at advancing the technology readiness level of morphing structures. Among these, Clean Sky [10] emerges as the most ambitious aeronautical research program ever undertaken in Europe, focusing on smart morphing structures. The primary goal of this program was to pioneer innovative technologies that could substantially improve the environmental performance of aircraft and air transportation, ultimately leading to quieter and more fuel-efficient airplanes. In line with this objective, a novel full-scale prototype of a morphing flap was successfully engineered for potential integration into the next generation of eco-friendly regional aircraft. Initially, research efforts concentrated on a specific segment of the flap component as the foundational stage of this advancement.

The initial phase of the research involved conducting computational fluid dynamic analysis (CFD), which demonstrated the feasibility of achieving higher maximum lift coefficient (Cl_max) values which is one of the important parameters in the assessment of aircraft performance and an increased stall angle through the active control of the flap’s camber, even with relatively small flap deployment angles [24,25,26]. Remarkably, the results indicated that a conventional double-slotted flap could potentially be replaced by a single-slotted flap with camber morphing capabilities while maintaining comparable effectiveness.

To guide aerodynamic investigations towards viable structural alternatives, a morphable design for the flap ribs was initially identified. A segmented arrangement, resembling fingers, was considered as a physical method to transition from the original airfoil shape to the desired configurations [27,28]. In Figure 1 [29,30], each rib was conceptualized as comprising four sequential blocks (B0, B1, B2, and B3), connected by hinges along the airfoil camber line (A, B, and C). Notably, Block B0 remained securely anchored to the flap track mechanism, while the remaining blocks could pivot around the hinges on the camber line. This unique arrangement effectively transformed the camber line into a flexible chain of successive segments. The introduction of linking rod elements and rotation amplification mechanisms (L, M1, and M2), strategically hinged on non-adjacent blocks, facilitated controlled rotation of the camber line segments, using specific gear ratios to achieve the desired shapes.

These components created a one-degree-of-freedom (1-DOF) system for each rib: preventing the rotation of any block restricts the ability to alter the shape, while activating an actuator to move any block results in a synchronized motion across all other blocks. The rib kinematics were transferred to the overall trailing-edge structure using a multi-box layout (Figure 2). In this structural design, each box featured a single-cell configuration defined along the span by corresponding blocks from consecutive ribs and along the chord by longitudinal stiffening elements, including spars or stringers. Four independently controlled rotary actuators were installed within the box to enable the morphing capabilities.

2. Motivation of This Study

Building on the promising results of the Clean Sky program, a new initiative was launched to further investigate multipurpose adaptive structures. Clean Sky 2, the continuation of Europe’s largest aviation research and innovation effort, was formally inaugurated in 2014 by the European Commission to meet societal expectations and compete globally by developing efficient and environmentally friendly aircraft solutions. A key aspect of the Clean Sky 2 project is establishing a strong collaboration with the European Aviation Safety Agency (EASA), formalized through a Memorandum of Cooperation (MoC). EASA is responsible for the potential certification of Clean Sky technologies. This collaboration includes various components, including mitigating risks, demonstrating the viability of new concepts and technologies introduced by the Clean Aviation initiative, advancing industry norms, developing innovative certification approaches and methods for aircraft and systems design, and aligning regulatory frameworks with other regulatory bodies and the International Civil Aviation Organization (ICAO) [31].

Clean Sky 2 hosts a wide range of innovative research initiatives that extend well beyond traditional industry examples in areas like cabin technology, production, maintenance, and hybrid propulsion. Among these initiatives, the Clean Sky 2—Airgreen 2 initiative investigated a groundbreaking multi-functional morphing wing technology aimed at enhancing the aerodynamic performance of next-generation turboprop regional aircraft throughout their entire flight trajectory.

The electromechanical configuration of the morphing-flap system is an appealing alternative in this research context. While smart materials technology has been applied to various shape-changing systems and has advanced the field, it remains unsuitable for aviation applications due to its high power consumption and need for a robust power system.

The primary goal was to optimize aircraft performance across various flight conditions, with a particular focus on reducing emissions during the program’s second phase. To achieve this, a groundbreaking system was developed to improve both high-lift capabilities and aerodynamic efficiency during the cruise phase of a standard 90-seat aircraft. This advancement was achieved through the innovative technique of in-flight camber morphing of the wing flaps. The newly introduced morphing design concept supports three different morphing modes, dynamically adapting to specific flight conditions.

Morphing mode 1: Camber morphing of the entire wing flap cross-section (Figure 3a).

Morphing mode 2: Deflection of +10°/−10° (upwards/downwards) at the flap tip segment, spanning from 90% to 100% of the local chord (Figure 3b).

Morphing mode 3: The flap tip is divided into three distinct tabs along the span. These tabs are differentially deflected both upwards and downwards during cruise, each within a range of +/−10°. This mode introduces aerodynamic twist and facilitates load control at higher speeds (Figure 3c).

The smart structure underwent a comprehensive design and validation process that included both ground and wind tunnel experiments, addressing the morphing modes required for low-speed (take-off/landing) and high-speed (cruise) conditions. The primary goal of the wind tunnel tests was to validate the CFD results and demonstrate that the scaled device performed as expected under simulated operating conditions. In contrast, the ground tests aimed to confirm that the full-scale system functioned as expected and was robust enough for flight preparation. For the ground tests, a true-scale device was manufactured, while the wind tunnel test article was scaled at 1:3 and tested at Mach numbers close to those anticipated in flight.

The wind tunnel model required a new morphing architecture since scaling the embedded mechanisms and actuators for shape transition was not feasible. This new design was developed to address the challenges of dimensional scaling, support the weights during aerodynamic testing, safely withstand the aerodynamic loads expected in the tests, and accurately replicate the shape transition capabilities of the full-scale flap.

The development of this sophisticated wind tunnel model was detailed in previous research [32], which explored the design processes and technological solutions used.

Initially, CFD studies were conducted to evaluate the most extreme load scenarios anticipated during the wind tunnel tests. This led to the development of a novel mechanical layout for the model, enabling rapid shape changes and minimizing dead time during flap aerodynamic configuration adjustments. To ensure safety, particularly under high-speed conditions, and to demonstrate the model’s robustness, in-depth structural evaluations were conducted [32].

A total of 18 test scenarios were generated using three Mach levels, two angles of attack, and three flap settings/shapes. A structural study was performed to evaluate the system’s safety in the wind tunnel, focusing on the two most important scenarios. The structural analysis, conducted using Ansys^® and based on pressure distributions from CFD, evaluated Von Mises stress distribution and maximum displacement across the flap. The results demonstrated that all components and joints had a significant safety margin corrected against plastic deformation and failure under the most severe load conditions expected during the tests [32].

In Figure 4, the project work packages are detailed, illustrating the progress made thus far. Sicim et al. explained the details of the design characteristics and CFD assessment of a novel morphing-flap system [32]. During the design phase, careful consideration was given to several factors, including ensuring that the solutions meet the standards of the aeronautics market, such as certification, safety, operational reliability, cost, and maintenance processes. Additionally, feasibility was demonstrated to meet the criteria of the aircraft manufacturing industry.

The focus of this study is on the work package highlighted within the red box, with a particular emphasis on the validation of the design and production processes for the wind tunnel test article. This discussion is complemented by experimental static tests conducted to verify the effectiveness of the structural and mechanical approaches implemented. It is crucial to validate structural analyses through testing to confirm the suitability of the morphing-flap structure for wind tunnel experiments, its ability to endure the aerodynamic forces encountered in the wind tunnel, and its capacity to maintain functionality under such conditions.

3. Mechanical Layout of the Morphing-Flap Wind Tunnel Model and Static Test Targets

The wing flap model, designed for wind tunnel validation, is scaled at 1:3, with relevant dimensions provided in

Table 1. Detailed dimensions of scale wind tunnel model.

Item	Dimension (mm)
Wing span	4466.23
Inboard flap span	1359.73
Outer flap span	1703.04
Inner flap root chord	292.77
Inner flap tip chord	292.77
Outer flap root chord	293.9
Outer flap tip chord	213.05

. Both the inner and outer flap models were developed using the same conceptual layout, carefully designed to withstand the most extreme operating load conditions anticipated during testing [32]. The overall structure was specifically defined for the inner segment, with the intention of applying a similar design approach to the outer segment.

The true-scale flap, as demonstrated by the wind tunnel model, successfully replicated the shape alterations of the flap in accordance with three distinct morphing modes. Morphing mode 1 involved comprehensive changes to the flap’s camber and was intended for actuation exclusively during take-off and landing phases (low-speed conditions). This mode improves high-lift performance, facilitating steeper initial climb and descent trajectories and contributing to noise reduction during these phases. It offers a wider range of airfoil configurations for each flap setting, potentially simplifying deployment mechanisms. This allows for the integration of actuation tracks within the wing’s airfoil, eliminating the need for external fairings. The morphing modes 2 and 3 relate to the final section of the flap in the chordwise direction and are designed to operate only during cruise conditions when the flap is stowed (Figure 3). These modes are aimed at improving the lift-to-drag ratio by facilitating load control functionalities.

Both the inner and outer flap models were designed to be manually adjusted according to the three morphing modes mentioned above. A consistent conceptual mechanical layout was applied to both flap segments, taking into account the most demanding operational conditions projected during the testing. The following description focuses on the general layout of the inner flap model, with the assumption that a similar configuration was implemented for the outer flap segment.

The inner flap model features a segmented design that mirrors the chordwise division of the full-scale device into four distinct blocks (S1, S2, S3, and S4, as shown in Figure 5). These blocks are hinged along three lines (H1, H2, and H3). The final block, S4, is further divided into three sections (Tab1, Tab2, and Tab3) along the span, each capable of independent rotation around H3 by equal or varying angles. The initial segment of the flap was securely attached to a series of brackets, which are bolted to the rear spar of the wing.

Rotations around the hinge axes from H1 to H3 are restricted by robust mechanical locks (can be seen in Figure 6, Item 3), which are firmly attached to the flap blocks.

All structural parts of the flap, including the flap brackets mounted on the wing’s rear spar, were fabricated using numerical control milling from large metal blocks. The materials used for this operation are listed in

Table 2. List of materials used in each part of the flap.

Bill of Material
Item Number	Part Description	Material
1	Flap Brackets	Al7075–T6
2	Upper Covers	Al2024–T351
3	Hinges	Uddeholm Ramax Steel®
4	Lower Covers	Al7075–T6
5	Lockers	Al7075–T6
6	Tabs	Al7075–T6

.

The structural layout was defined based on numerical stress analysis, with detailed results described in [32]. This study examined the most severe load combinations of flap configurations and expected test loads (derived from CFD analysis) to carefully select materials, dimensions of the structural components, and all joints. The test article was numerically verified to comply with wind tunnel safety requirements, maintaining a safety margin of three against local plasticization even under the most severe load conditions.

To experimentally confirm these findings, a static test was conducted on the manufactured test article before its installation in the wind tunnel. Due to budget and time constraints, the static-test process was simplified and did not exactly replicate the most severe load distribution expected during testing. Instead, three reference stations along the span of the flap were arbitrarily selected, and the resultant of the highest pressure distribution (150 kg) was applied using a simple wiffle tree loaded by a single hydraulic jack. The spacing of the sections was chosen to minimize the cost of the wiffle tree.

The simplified load distribution was numerically replicated to validate the finite element model of the test article and confirm the reliability of its predictions under simulated pressure distributions. Given the expected elastic displacements and strains and the acceptable aerodynamic impact of any deviations from their theoretical values, the model was validated if the predicted displacements and strains for the simplified load conditions aligned with experimental outcomes within a margin of error of ±10% for displacements and ±5% for strains.

4. Structural Analysis of the Morphing-Flap Wind Tunnel Model

A thorough investigation into the flap test object was undertaken to assess the viability of the proposed mechanical solutions and ensure the safety of the testing procedures. Utilizing the Ansys Workbench^® 2023 R1 platform, linear static analyses were performed through the finite element (FE) method. This analysis focused on understanding the stress distributions in the model’s critical components under extreme loading conditions. The design’s reliability was validated by demonstrating the absence of permanent deformations and localized failures, along with confirming sufficient safety margins. Following the finite element analysis, stress-hand calculations were executed to evaluate the structural soundness of all joints connecting the various components of the model. The primary focus was on the inboard flap, given its structural similarity to the outer section and the anticipated high loads during wind tunnel tests. It was assumed that all flap components, including the brackets affixed to the rear spar of the wing, would be crafted using numerical control machining from large metallic blocks. The relevant mechanical properties of the selected materials are detailed in

Table 3. Mechanical properties of materials used on FEA.

AL 7075-T6
Young’s modulus	E	71.70 GPa
Poisson ratio	v	0.33
Tensile strength yield	Fty	503 MPa
Tensile strength ultimate	Ftu	572 MPa
AL 2024-T351
Young’s modulus	E	73.1 GPa
Poisson ratio	v	0.33
Tensile strength yield	Fty	290 MPa
Tensile strength ultimate	Ftu	441 MPa
Uddeholm Ramax® Steel
Young’s modulus	E	215 GPa
Poisson ratio	v	0.28
Tensile strength yield	Fty	990 MPa
Tensile strength ultimate	Ftu	1140 MPa

.

Tetrahedral elements, appropriately sized to capture the stress concentration factors surrounding holes and fillets, discretize the lower and upper covers of the initial three blocks (B1–B3). Conversely, hexahedral elements mesh flap brackets, lockers, hinges, and tabs were used because of their superior performance in terms of convergence rates and solution accuracy. The geometric complexity of these components exceeds that of the flap covers, thereby increasing the likelihood of failure in regions characterized by high curvature. The pins and bolts have been modeled using beam elements, with remote grids employed to connect the center of the bolt to the contact area with the bolt head or nut.

Table 4. Mesh data (number of nodes and elements).

Mesh Details
Element Type	Area	Number of Elements
Tetra	Upper Covers, Lower Covers, Flap Brackets	1,121,761
Hex Elements	Tabs, Lockers, Hinges, Flap Brackets	685,350
Beam Elements	Bolts, Pins	153

summarizes the mesh data, detailing the number of nodes and elements utilized. The mesh consists of 1,121,761 tetrahedral elements, 685,350 hexahedral elements, and 153 beam elements. To accurately simulate the fasteners, beam elements were strategically positioned at the locations of the pins and bolts, and the mesh was refined in areas identified as having the potential for maximum stress concentration. The integration of the upper and lower covers of the flap and brackets was achieved through the use of tetrahedral elements, while hexahedral elements were utilized for the other components. Figure 7 depicts the detailed mesh generation across the different sections.

The quality of the mesh significantly influences the outcomes derived from finite element analysis (FEA). A carefully constructed mesh ensures that the physical phenomena being simulated are accurately represented, thereby yielding reliable and precise results. High-quality meshes are proficient in capturing critical stress concentrations and effectively modeling the material’s behavior under various loading conditions. A converged mesh, in terms of numerical accuracy, is characterized by the absence of significant differences in results when mesh refinement is applied. Mesh convergence is particularly important when developing a model aimed at capturing peak stress or strain. Consequently, it is essential to conduct a convergence study in regions of peak stress to confirm that the mesh size is sufficiently refined to accurately represent the phenomena of interest and the critical stress.

Enhancing the number of nodes contributes to improved accuracy of results; however, it simultaneously results in an increase in both solution time and associated costs. A prevalent strategy involves augmenting the number of elements specifically in regions subjected to high stress, rather than uniformly reducing the global element size throughout the entire model. This process is sustained until the variation between two successive results is less than 5%.

Figure 8 presents the outcomes of four distinct experimental runs. It displays the variation in stress value in the critical region according to the number of meshes. It was observed that the discrepancy between the outcomes of the third and fourth cases was 2.8%. As a result, the mesh convergence analyses were concluded after the fourth iteration, and the subsequent sections of this study were carried out using the mesh configuration established during the fourth iteration.

A force of 50 kg was applied at each of the three locations determined in a static test. The load was restricted to the joints of the wing spar. The contact between the connecting surfaces of the lockers, hinges, and blocks is modeled as frictionless. These components were assembled using bolts, which are classified as beam elements. Additionally, a similar relationship was established between block S1 and the flap brackets. The entire model is constrained in all three translational degrees of freedom due to the flap’s connection to the wing at the rear surface of the brackets.

5. Selection of Strain Gauge Locations for the Static Test of the Model

Strain gauges were placed along the areas where the maximum strain was expected on the base of numerical structural analyses. Four different areas for the outer flap and three different areas for the inner flap were identified. The regions designated for the outer and inner flaps and the associated strain distributions are shared in Figure 9 and Figure 10. Facing obstacles when attempting to precisely position strain gauges in areas of peak stress was expected due to geometric variances such as curvature or inconsistencies in thickness. In this research, efforts were made to locate the gauges as proximate to these points as feasible. Also, it was observed that not all detected areas were physically suitable for attaching strain gauges on the flap.

Therefore, three strain gauges have been attached to the outer flap and three strain gauges to the inner flap.

Strain gauge number 1 was mounted on the first locker to the left. Meanwhile, strain gauge number 2 was positioned immediately beside the first locker on the right-hand side and strain gauge number 3 was mounted on the rear outer surface of the S1 segment. For a detailed visual representation of where these strain gauges were placed, one can refer to Figure 11, which provides a comprehensive illustration of their locations.

Strain gauges were attached to the inner flap in locations as close as possible to the high-strain areas indicated in Figure 10. As expected, Figure 10 shows that the areas near the flap brackets have the highest strain concentrations. Nevertheless, due to the curvature and unsuitability of these areas for strain gauge attachment, points 1, 2, and 3 were selected. Figure 12 shows the strain gauges that were adhered to the inner flap. Strain gauge 1 was positioned slightly below the flap bracket on the lower cover surface of segment 1. Strain gauge 2 was adhered to the lower cover surface of segment 2. Strain gauge 3 was mounted on the middle locker, which connects segment 1 with segment 2.

6. Static-Test Setup: Tools and Measuring Equipment

The primary objective of conducting a static test was to validate the flap’s capability to withstand limit loads without permanent distortions, failures, or structural buckling. To ensure the accuracy and reliability of the results, meticulous numerical simulations were executed, serving several purposes:

• Approach the experiment campaign in a logical manner with a focus on promptly identifying any potentially hazardous deviations between the actual structural behavior of the flap and its expected behavior.
• Precise determination of the optimal location on the prototype for installing deformation and displacement sensors was carried out, guaranteeing comprehensive data collection during the test.
• Rigorous verification of the adequacy and effectiveness of the test rig and the load transfer mechanism was conducted, ensuring their suitability for accurately simulating real-world conditions.

To achieve these objectives, a comprehensive finite element model, previously utilized during the flap design process, was merged with the finite element model of the test rig. This integration enabled the accurate reproduction of load transmission paths, commonly referred to as the “whiffle tree”, and the constrained conditions that are expected during the tests.

Figure 13 shows the assembled model, which combines the flap and the test rig. Figure 13a–c show the isometric, side, and front views of the test rig, respectively, while Figure 13d shows the method of attaching the flap to the test rig. The loading arrangement depicted in Figure 13 utilized the whiffle tree concept. Testing was conducted using inner and outer flap blades. The whiffle tree allowed for multiple loads to be applied using a single-load introduction. This loading approach offers the advantage of requiring fewer hydraulic actuators compared to conventional methods involving multiple loading or a single crank. To emulate the hydraulic jack’s action located at the top of the whiffle tree, a single consolidated force of 150 kg was applied. A force of 150 kg has been chosen for application, as it represents the maximum capacity of the linear actuator.

On the basis of pre-test analyses, strain gauges were positioned in the most stressed zones of the rib links and spar, being the maximum expected strain sensor (MSS) located on the second rib linking beam element.

The 150 kg traction load was divided into three loads of 50 kg through the load distribution frame. It is shown with a purple arrow in Figure 14.

Load data have been measured by the use of a load cell; displacements have been measured by the use of two linear transducers and one rotative transducer (potentiometers); deformations have been measured by the use of three couples of strain gauges positioned on the base of numerical stress analysis. Figure 14 shows the overall test setup and identifies the different components.

The instrument specifications are reported in

Table 5. List of instruments.

Item	Model	Measurement Error (%)
Lms Scadas	Daq Mobile (Siemens, Leuven, Belgium)	-
Load Cell	TS—200 kg Class C2 (Vetek, Väddö, Sweden)	0.03
Linear Actuator	L11TGF12V50-T-1 (Eco-Worthy, Xiamen, China)	-
Rotative Displacement Transducer	Enosis Sensor (Enosis Electric, Singapore)	0.02
Linear Displacement Transducer	Gefran PA-12-A-50 (Gefran SPA, Brescia, Italy)	0.01
Strain Gauge	Mono axial1-LY13-3/120 (Ensinger, Nufringen, Germany)	-

.

For the application of load via the frame onto the flap, it was essential to bond the frame onto the upper side of the flap from these specific surfaces illustrated in Figure 14. To assure the adhesion remains effective throughout the testing process, it was necessary to perform an adhesive strength test.

Three different adhesives listed in

Table 6. Adhesive list.

Number	Item	Test Results	Max Applied Load (kg)
1	X60 Cold Curing Glue (Hottinger Baldwin Messtechnik, Darmstadt, Germany)	Pass	100
2	Pattex Millechiodi Forte & Rapido (Pattex, Milano, Italy)	Fail	45
3	3M Scotch-Weld Epoxy Adhesive 2216 (3M Company, Minnesota, USA)	Pass	100

were tested. The surfaces applied with adhesives 1 and 3 successfully sustained loads up to an approximate 100 kg, respectively, and thus met the testing criteria. However, the surface to which adhesive 2 was applied failed when it detached at a load of 45 kg. Despite adhesive 1’s rapid curing time and ease of application, it was difficult to it remove from surfaces without leaving residue after curing, leading to the decision to use adhesive number 3 during structural testing. Figure 15 shows the adhesive test setup.

7. Test Procedure

Step 1: The inner flap was installed onto the support apparatus. The flap, as depicted in Figure 16, was rigidly attached to the support frame.

Step 2: Two linear displacement transducers were positioned at each end of the flap, while a rotary displacement transducer was placed at the center. The location of the transducers can be seen in Figure 17.

8. Results and Discussion

The morphing-flap technology has undergone a gradual and sustained maturation process, culminating in the development of a scaled prototype for final wind tunnel testing under realistic operational conditions. Building upon existing mature technological solutions that demonstrate functionality and robustness, new research avenues are being explored to further improve the morphing flap in relation to the diverse requirements associated with cost-effective and sustainable integration at the aircraft level. Consequently, a systematic re-engineering of the flap’s constituent elements is anticipated to facilitate the transition from a reliable integrated smart system to an economically viable and competitive aeronautical component. This section provides a comprehensive analysis and interpretation of the findings presented in this article. The experimental data are utilized to validate the corresponding finite element models, which are employed to predict structural behavior under diverse loading conditions. These data provide valuable insights into the interaction among the various integrated functionalities and the structural responses.

Figure 18 depicts the temporal evolution of the applied force exerted on the outer and inner flap during the testing phase. The methodology for the load application involves meticulous increments in adjustments to ensure a steady progression towards the desired static load level across the flaps. Once the criterion for the static load is satisfied, the procedure transitions to an unloading phase, which systematically reduces the applied force to assess the structural response during load reversal.

Figure 19 displays the displacement’s evolution as a function of time. The findings highlight two key points. Firstly, the outer flap mechanism reaches a displacement apex of 15.78 mm under the maximal load conditions. Under the same load conditions, the maximum displacement of the outer flap mechanism is 6.77 mm. Secondly, the change in displacement demonstrates remarkable consistency across the range of loads applied. The measurement error, associated with the applied load, was determined by calculating the standard deviation of the measured displacements at the selected locations during the experiment.

Figure 20 shows the time-dependent variation in strain as measured by the strain gauges depicted in Figure 10 and Figure 12.

The strain measurements were captured using data acquisition software. It is unfortunate that the sensor designated for the measurement of e3x experienced a failure, and notably, a parallel malfunction was observed in the sensor for the outer flap. Nonetheless, this particular problem is not critical for the FE model’s validation process. It has been observed that the strain values change in correlation with the applied load and the observed displacement.

The measured displacement and strain values for the inner flap with those obtained from FEA are compared in

Table 7. Test and analysis results for inner flap.

Location	Test Results	Analysis Results	Ref. Figure	Error Rate %
d1 (mm)	6.47	6.24	Figure 21a	3.55
d2 (mm)	6.39	6.18	Figure 21a	3.28
d3 (mm)	6.77	6.25	Figure 21a	7.68
e1x (ms)	47.81	42.77	Figure 21b	10.54
e1y (ms)	37.96	37.96	Figure 21b	0.02
e2x (ms)	75.41	77.89	Figure 21c	−3.38
e2y (ms)	50.67	49.57	Figure 21c	2.18
e3x (ms)	-	85.34	-	-
e3y (ms)	61.69	62.35	Figure 21d	−1.06

. In Figure 21a, the changes in displacement are illustrated, with the variations occurring at points d1, d2, and d3 clearly marked.

d1 represents the data from the linear displacement transducer positioned at the right trailing edge of the flap, illustrated in Figure 14. d2 corresponds to the measurements from the linear displacement transducer at the left trailing edge of the flap. Furthermore, d3 includes data from the rotary displacement transducer at the midpoint of the flap’s trailing edge. The maximum displacements measured for d1, d2, and d3 are 6.47, 6.39, and 6.77, respectively. The error rates of the FE model analysis results were 3.55%, 3.28%, and 7.68%, respectively.

The strain gauge intended for measuring the e3x value was damaged during the experiment; therefore, its data are not available. For the remaining strain gauges, the maximum magnitudes of normal strain in the x and y directions were approximately 85 and 62 microstrains, respectively. Both peaks occurred at the location marked in Figure 12c, situated in the second locker, which houses the most load-bearing parts. This high strain is expected due to additional stress from load transitions between blacks and the moments induced by test loads on the lockers.

The measurement inaccuracies for normal strains were evaluated, revealing an average error of 2.5 microstrains ± 5 microstrains.

Figure 21 and the references in Table 7 exhibit the distributions of displacement and strain in the specified strain gauge regions obtained from the FE analysis.

Table 8. Test and analysis results for outer flap.

Location	Test Results	Analysis Results	Ref. Figure	Error Rate %
d1 (mm)	9.62	10.91	Figure 22a	−13.4
d2 (mm)	15.78	14.35	Figure 22a	9.06
d3 (mm)	14.14	14.03	Figure 22a	0.77
e1x (ms)	112.02	108.2	Figure 22b	3.41
e1y (ms)	124.827	130.4	Figure 22b	−4.46
e2x (ms)	38.15	36.74	Figure 22c	3.69
e2y (ms)	49.833	48.88	Figure 22c	1.91
e3x (ms)	-	159.39	-	-
e3y (ms)	207.594	203.65	Figure 22d	1.89

presents the comparison of displacement and strain for the outer flap, including data from both physical measurements and FEA. The maximum displacements labeled as d1, d2, and d3, were 9.62 mm, 15.78 mm, and 14.14 mm, respectively. The corresponding error rates in the finite element model analysis were 13.4%, 9.06%, and 0.77%. The strain gauge intended for measuring the e3x strain was damaged during testing; hence, its data are not available. For the remaining strain gauges, the maximum normal strains under the full load were approximately 203 microstrains in the x direction and 159 microstrains in the y direction. These peak strain values were observed at the location indicated in Figure 10, on the back of the flap. Increased strain in this area is anticipated due to the stress on the inner locker from the load’s reaction force. An evaluation of the normal strain measurement errors yielded an average error of 3.5 microstrains, with a standard deviation of ±6.5 microstrains.

The visuals in Figure 22, mentioned as references in Table 8, display the patterns of displacement and strain in areas measured by the strain gauges, as determined by the FE analysis. An illustration of the changes in displacement can be seen in Figure 22a. The alterations that take place at positions d1, d2, and d3 are clearly highlighted.

9. Conclusions

This article evaluates the design, structural analysis, and structural testing of the morphing-flap concept. The experimental work enhances understanding of the structural integrity of the morphing-flap prototype and validates the design methodology used for its fabrication.

The test article’s structural layout was determined through numerical stress analysis, which required examining extreme combinations of flap configurations and test loads. The article has been confirmed to meet wind tunnel safety standards, with a safety margin of three to account for local plasticization under extreme load conditions. To validate these findings, a stationary test was carried out prior to the wind tunnel installation. Due to resources and time constraints, the testing procedure was simplified, avoiding an exact replication of the most severe load distribution. Three reference stations were randomly chosen along the flap’s span, and a maximum pressure distribution of 150 kg was applied using a basic wiffle tree. This simplified load distribution was numerically replicated to validate the finite element model and verify its reliability under simulated pressures in the wind tunnel. The model’s validation was determined by comparing the anticipated displacements and strains to the experimental outcomes, with an acceptable tolerance of ±10% for displacements and ±5% for strains.

The only significant divergence from the target values was observed for displacements d1. Despite this, the model was deemed “validated” because the variation was attributed to the elasticity of the test article constraint, which was not replicated in the FE model where all constraints are assumed to be perfectly rigid. Hence, the displacement at the edge of the flap during the testing slightly exceeded the expected value.

The comparison of curves derived from physical experiments with those generated by mathematical models is a critical methodology employed by researchers to evaluate the adequacy of a model in representing physical phenomena. This approach facilitates the validation of simulation models, particularly in scenarios where only a limited set of experimental data is available, thus providing a distinct advantage over alternative validation techniques. The validated finite element analysis (FEA) model plays a crucial role in advancing the effective development of new designs that will emerge during the comprehensive design of morphing flaps. Furthermore, the application of this validated model is expected to result in reductions in both costs and time associated with the development of new designs.

The discrepancy is primarily due to the flexibility of the test substance restriction. Nonetheless, the fact that the results remain within acceptable limits reinforce the validity of the FE model. Considering the elasticity of the test article and difference between experimental and FEM results, the notable consistency between values from measurements and those from computational results validates the approach employed in constructing the finite element (FE) model.

This led to the conclusion that the modeling methodologies used throughout the testing and design stages were entirely reliable. The results inferred from calculations under ultimate load levels, relevant to the wind tunnel test campaign, further supported the suitability of the scaled flap prototype for testing.