Design, Manufacturing and Experimental Validation of an Integrated Wing Ice Protection System in a Hybrid Laminar Flow Control Leading Edge Demonstrator

Abstract

This paper presents the design, manufacturing, instrumentation and validation by tests (ground and icing wind tunnel) of a full-scale Hybrid Laminar Flow Control (HLFC) leading-edge demonstrator based on Airbus A330 outer wing plan-form. The Ground-Based Demonstrator (GBD) was developed to reproduce a full-scale, realistic wing section integrating into the leading-edge three key systems: micro-perforated skin for the hybrid laminar flow control suction system (HLFC), the hot-air Wing Ice Protection System (WIPS) and a folding “bull nose” Krueger high-lift device. The demonstrator combines a superplastic-formed and diffusion-bonded (SPF/DB) perforated titanium skin mounted on aluminum ribs jointed with a carbon-fiber-reinforced polymer (CFRP) wing box. Titanium internal ducts were designed to ensure uniform hot-air distribution and structural compatibility with composite components. Manufacturing employed advanced aeronautical processes and precision assembly under INCAS coordination. Ground tests were performed using a dedicated hot-air and vacuum rig delivering up to 200 °C and 1.6 bar, thermocouples and pressure sensors. The results confirmed uniform heating (±2 °C deviation) and stable operation of the WIPS without structural distortion. Relevant tests were performed in the CIRA Icing Wind Tunnel facility, verifying the anti-ice protection system and Krueger device. The successful design, fabrication, testing and validation of this multifunctional leading edge—featuring integrated HLFC, WIPS and Krueger systems—demonstrates the readiness of the concept for subsequent aerodynamic testing.

1. Introduction

The continuous growth of air traffic and increasingly stringent environmental targets imposed by ACARE and ICAO [1] have intensified the search for technologies capable of significantly reducing fuel burn and emissions for future transport aircraft [2].

Multiple technological pathways have been pursued to reduce fuel consumption in commercial aviation, including improvements in propulsive efficiency and the adoption of alternative fuels, the integration of advanced lightweight materials and optimized aerodynamic configurations, as well as the implementation of flow-control systems aimed at reducing aerodynamic drag [3,4]. In this context, laminar-flow technologies—and, in particular, Hybrid Laminar Flow Control (HLFC)—have emerged as a promising solution capable of delivering substantial fuel-burn reductions by lowering skin-friction over lifting surfaces [5].

HLFC offers the potential for friction-drag reductions that translate into fuel savings on long-range missions [6]. HLFC combines a relatively small amount of active boundary-layer suction near the leading edge with a downstream airfoil pressure distribution tailored for natural laminar flow (NLF), thereby relaxing some of the stringent tolerances associated with purely passive NLF wings while maintaining attractive performance gains [7].

Over the past decades, extensive numerical and experimental work has been devoted to assessing HLFC concepts on generic airfoils, swept wings and complete aircraft configurations. Early feasibility studies and design investigations at NASA and DLR already highlighted the strong potential of laminarization for transonic transport aircraft [8,9,10], while more recent work has focused on realistic wing architectures, including the integration of suction systems, structural constraints and system-level aspects [6,7,8,9]. For example, Jentys et al. performed high-fidelity CFD analyses of a transonic transport wing equipped with HLFC, showing substantial reductions in drag and fuel burn compared to a turbulent baseline [5], whereas Karpuk et al. investigated HLFC capabilities from an aero-structural perspective, quantifying performance benefits in the context of a medium-range commercial aircraft [11]. Complementary work at the Aircraft Research Association and within various European projects has addressed HLFC design methodologies, system architectures and experimental validations on wind-tunnel models and flight demonstrators [12,13,14].

A key enabler for practical HLFC implementation is the development of robust, manufacturable suction surfaces and internal flow circuits. Current solutions typically employ micro-perforated metallic or composite skins, combined with tailored inner chambering or, more recently, chamberless concepts relying on variable skin porosity [15,16,17,18]. Kilian and Horn successfully demonstrated a chamberless HLFC leading edge with an outer skin of variable porosity on a vertical tail-plane model at flight Reynolds number, using a CEAS Aeronautical Journal configuration tailored to industrialization and weight reduction [17]. In parallel, studies have examined the structural behavior and manufacturability of Ti-6Al-4V skins for HLFC applications [16], as well as the integration of suction panels into wings for all-electric short-range aircraft [15]. Additional research at DLR has explored additive and laminated porous sheets as alternatives to etched or laser-drilled panels, aiming to extend the design space and reduce manufacturing effort while preserving the required pressure-loss characteristics [18].

Despite these advances, maintaining laminar flow in operational conditions remains challenging due to the extreme sensitivity of boundary-layer stability to surface contamination, roughness and degradations. Even small perturbations caused by insect residue, erosion or surface waviness can trigger premature transition, thereby reducing HLFC benefits [6,19]. Ice accretion is particularly critical for laminar wings because it modifies both the geometry and roughness of the leading edge, often leading to large performance penalties or even loss of controllability. Comprehensive reviews by Bragg et al. and Gent et al. have documented the flow field physics and aerodynamic consequences of various ice shapes—including roughness, horn ice and spanwise ridges—as well as the performance degradation mechanisms on iced airfoils [20]. NASA and other research organizations have subsequently developed advanced numerical tools and wind-tunnel facilities to support the design and certification of aircraft ice protection systems (IPSs) [21].

For Hybrid Laminar Flow Control (HLFC) wings, icing imposes stringent requirements on the Wing Ice Protection System (WIPS), which must be integrated directly beneath the micro-perforated suction surface. Conventional thermal anti-icing solutions, based on hot-air or electro-thermal systems, must operate within a highly constrained volume while ensuring uniform heating and avoiding thermal gradients or structural distortions that could degrade suction uniformity and laminar-flow stability [16,17,22]. Moreover, WIPS components must be compatible with hybrid structures combining titanium leading edges, CFRP wing boxes and aluminum sub-structures, resulting in a strong coupling between aerodynamic, thermal and structural design. Although recent studies have explored advanced HLFC leading-edge concepts, including chamberless and variable-porosity configurations, fully integrated ice-protected HLFC wings remain at a relatively low Technology Readiness Level (TRL) [13,14,18,23].

In parallel, the integration of an efficient high-lift system is essential for protecting the laminarized leading edge during take-off and landing. Krueger flaps have emerged as a promising alternative to conventional slats for laminar wings, providing both lift augmentation and protection against surface contamination [24,25,26]. While experimental and numerical studies have demonstrated their aerodynamic viability, the combined integration of ice protection and Krueger systems within HLFC wings remains a significant challenge. Large European programmes such as AFLoNext and Clean Sky have therefore focused on integrated demonstrators combining HLFC, ice protection and high-lift systems [13,14], yet comprehensive full-scale validations of such multifunctional leading edges are still scarce in the open literature.

This article presents selected results that focused on the development and validation of several large-scale demonstrators implementing advanced flow-control technologies. Within this context, this study examines the integration of a Hybrid Laminar Flow Control (HLFC) system—based on a micro-perforated titanium skin and incorporating an anti-icing protection solution into a full-scale wing segment demonstrator equipped with a “bull-nose” Krueger configuration. Particular emphasis is placed on the manufacturing process and the experimental validation of suction and thermal performance. This study contributes to the development of an integrated solution capable of ensuring both effective ice protection and the preservation of laminar-flow control—an essential requirement for the application of HLFC technology in next-generation commercial aircraft.

2. Conceptual Design and Functional Requirements

The present study focuses on a representative wing segment of the Airbus A330 aircraft. A baseline HLFC (Hybrid Laminar Flow Control) panel was provided by Airbus, derived from the outer wing platform of the A330 model and corresponding to two design sections, as illustrated in Figure 1.

The primary objective of the Ground-Based Demonstrator (GBD) development was to reproduce, at full scale, a leading-edge wing section equipped with fully functional systems for laminar flow suction (HLFC), anti-icing protection (WIPS—Wing Ice Protection System) and a Krueger-type high-lift device. This demonstrator served as a validation platform for assessing multi-material integration and the simultaneous operation of aerodynamic, thermal and structural subsystems. The GBD was specifically designed to evaluate the compatibility between a CFRP (Carbon Fiber Reinforced Polymer) wing box structure and a metallic leading-edge section incorporating active HLFC and WIPS technologies. Geometrically, the demonstrator represents an outer wing portion with a 2.3-m span, comprising the wing box, the fixed leading edge (FLE) and a movable Krueger panel (Figure 2).

The overall assembly adopts a hybrid material concept: the wing box is manufactured from CFRP composite, while the leading-edge section and the HLFC/WIPS components employ titanium and aluminum alloys to ensure adequate thermal and mechanical resistance. Through the integration of these subsystems, this study pursued three major objectives:

✓

To demonstrate the technological feasibility of manufacturing and assembling the titanium perforated skin using the SPF/DB (Superplastic Forming and Diffusion Bonding) process;

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To achieve the functional integration of the suction and anti-icing systems within a compact, shared and easily accessible architecture;

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To validate the kinematic behavior of the Krueger panel and its structural compatibility with the laminar leading-edge configuration.

This work therefore provides an essential experimental foundation for future applications of integrated Hybrid Laminar Flow Control and ice protection technologies in next-generation commercial aircraft, where aerodynamic efficiency, system integration and structural reliability must coexist within stringent design constraints.

The leading-edge structural assembly was designed to integrate components with distinct mechanical and thermal properties into a coherent and functional configuration. It comprises metallic ribs made of 7050-T7651 aluminum alloy, precisely machined to ensure effective load transfer between the skin and the spar. A bird-strike protection sub-spar, manufactured from 7475-T761 aluminum alloy with a thickness of 4.5 mm, provides additional structural integrity in the event of impact. The assembly also incorporates front landing brackets and lower fixation straps (SS Straps) that ensure proper attachment and stability of the leading-edge elements. In addition, titanium thermal shields are installed to provide insulation for the regions exposed to elevated temperatures generated during the operation of the Wing Ice Protection System (WIPS) (Figure 3).

The Hybrid Laminar Flow Control (HLFC) and Wing Ice Protection System (WIPS) were integrated within a unified leading-edge architecture designed to combine, in an efficient and compact manner, the functionalities of laminar flow control and thermal protection. The configuration was conceived to ensure structural, thermal and functional compatibility between the metallic and composite subsystems forming the leading-edge assembly.

The final design comprised two primary flow circuits: the first dedicated to hot-air distribution for anti-icing purposes and the second to air suction through the perforated titanium skin. Both networks were interconnected via a common control valve capable of simultaneously regulating pressure and mass flow rate in the two systems. This integrated dual-line solution reduced overall system mass and complexity, while maintaining the reliability and safety standards required for aerospace applications.

All ducts, manifolds and fittings were manufactured from Ti-6Al-4V titanium alloy, selected for its high mechanical strength at elevated temperatures and compatibility with dry suction environments. The pneumatic lines were thermally insulated to prevent mutual interference between the hot and cold zones of the structure, while the routing geometry was optimized through 3D CAD analyses to preserve the clearance necessary for the Krueger panel kinematics.

The internal configuration of the perforated titanium skin chambers was finalized following an extensive geometric optimization process. The dimensional requirement drawing was subsequently approved, freezing the design and defining the critical HLFC parameters, including suction pressure uniformity, distribution channel length and manifold port positioning.

To validate the installation envelope and mechanical interfaces between the HLFC and WIPS subsystems, a full-scale 3D-printed dummy valve was fabricated. This mock-up accurately reproduced the physical volume of the functional components, enabling direct assessment of assembly accessibility, maintainability and inspection feasibility. Concurrently, the physical verification confirmed compliance with segregation criteria between pneumatic circuits and electrical control zones, thereby ensuring system integrability within the leading-edge environment.

Figure 4 presents the final integrated configuration of the HLFC and WIPS subsystems, highlighting the arrangement of the internal suction chambers, hot and cold air manifolds, main fittings and the control valve implemented in the geometric simulations. Detailed renderings of the ducting and connecting components, including the 3D-printed dummy valve, emphasize the high level of integration and design sophistication achieved between the aerodynamic and thermal functionalities of the system.

This stage allowed for a thorough validation of the functional interfaces between the wing structure, the pneumatic circuits and the actuation mechanisms. The results confirmed that these subsystems can be successfully integrated within a laminar-flow aerodynamic configuration without introducing penalties in terms of maintainability or operational safety.

In the HLFC demonstrator, the Krueger panel was implemented as a deployable element positioned beneath the leading edge. Its design serves two primary functions: to increase lift during low-speed flight and to protect the laminarized surface during cruise conditions. The concept was developed to reproduce the behavior of a full-scale high-lift device while maintaining strict geometric and functional compatibility with the laminar leading-edge configuration.

The Krueger system for the HLFC ground-based demonstrator [27] was designed in accordance with a set of geometric and aerodynamic constraints defined to ensure compatibility with laminar-flow operation (Figure 5). Specifically, the Krueger configuration was required to exhibit a minimum chord, measured normal to the leading edge, equal to 11% of the sectional reference chord. For the outboard HLFC panel, which features a reference chord of 3.489 m, this requirement yields a minimum normal-to-LE Krueger chord of 0.349 m.

In the stowed position, the forward extent of the Krueger panel was mandated to remain no farther forward than 1% of the reference chord and to lie aft of the cruise aerodynamic attachment line. This ensures that the stowed high-lift device does not interfere with the suction-based laminarization strategy or disturb the pressure distribution over the leading edge during cruise.

Furthermore, the Krueger aerodynamic surface and associated kinematic mechanisms were required to preserve the ability to apply laminar-flow suction across the full span of the ground-based demonstrator panel, except for small margins near the panel edges where spanwise joints are expected to generate a localized turbulent wedge (Figure 6). All structural attachments linking the Krueger system to the HLFC leading edge were incorporated directly into the demonstrator to provide a representative integration environment.

The finalized Krueger flap configuration enables a rotation of approximately 140° between the fully deployed and fully retracted positions, achieved through a goose-neck–type kinematic mechanism that satisfies all geometric and operational constraints related to hinge-point placement and actuator integration. The bull-nose element undergoes an additional rotation of about 125°, which substantially improves the available stowage volume in the retracted position. The kinematic layout is further characterized by a delayed deployment of the bull-nose during the overall deflection sequence, thereby reducing transient aerodynamic loads acting on the mechanism.

Because the kinematic assembly intersects the lower surface of the wing, dedicated cut-outs were incorporated into the lower Leading Edge structure. These openings are sealed in cruise conditions by rigid sealing plates attached to both the flap and the kinematic links. The foremost points of the sealing plates and corresponding cut-outs were verified to remain downstream of the attachment line, ensuring that the laminar flow on the upper wing surface remains undisturbed during cruise.

A final, critical feature of the designed mechanism is that, in its vertical orientation, the Krueger flap remains fully below the wing leading-edge contour. This prevents unintended shielding of the wing flow and preserves the aerodynamic functionality of the leading edge throughout the operational envelope.

The assembly consists of the actuator lever, the drive link, the eccentric rigging device for fine adjustment and the gooseneck-type support brackets that attach the mechanism to the internal structure (Figure 7).

Finite Element Method (FEM) analyses performed in NASTRAN and PATRAN confirmed the mechanical integrity of the mechanism, indicating safety factors between 1.22 and 1.29 under maximum load conditions, including bird-strike impact scenarios. To provide overload protection, a calibrated shear pin was incorporated at the interface between the drive shaft and the main lever, designed to fail at a load of 2.3 ± 0.1 kN, thereby preventing damage to the adjoining joints.

The wing box, Figure 8, constitutes the primary load-bearing element of the demonstrator, integrating both composite and metallic components within a hybrid architecture optimized for stiffness, weight efficiency and manufacturability. The main spar and the upper and lower covers are made of carbon-fiber-reinforced polymer (CFRP) to ensure a high strength-to-weight ratio, whereas the internal ribs are metallic, providing improved load transfer capability and damage tolerance.

Two types of ribs were employed in the design: end ribs, machined from 7050-T7651 aluminum alloy, which provide structural reinforcement and mounting interfaces for subassemblies and intermediate ribs made of CFRP laminates. To compensate for thickness variations inherent to composite laminates, liquid shims were applied to guarantee uniform contact between the ribs and the skin panels. The joints were secured using blind bolts, ensuring a smooth external surface consistent with aerodynamic requirements.

In the leading-edge region, titanium thermal shields were installed to thermally isolate the composite spar from the Wing Ice Protection System (WIPS) ducting, thus preventing thermal degradation during anti-icing operation.

The entire assembly was designed with modular architecture, allowing rapid disassembly and inspection, a key requirement in the aeronautical field and a practical necessity during ground-based functional testing.

The Ground-Based Demonstrator (GBD) was conceived as an experimental platform for the development and validation of advanced technologies dedicated to Hybrid Laminar Flow Control (HLFC) at the leading edge of commercial aircraft wings. The primary objective of the demonstrator was to reproduce, at full scale, a representative wing section incorporating all relevant functional systems, thereby enabling integrated testing of multi-material architectures, movable mechanisms and pneumatic and thermal subsystems.

The demonstrator structure consists of a wing box (WB) manufactured from carbon-fiber-reinforced polymer (CFRP) and equipped with a main spar, upper and lower covers and a full-chord movable flap. The leading edge integrates titanium alloy components produced through SPF/DB (Superplastic Forming and Diffusion Bonding) technology, along with the Wing Ice Protection System (WIPS), the HLFC suction network and the movable CFRP Krueger panel with its actuation mechanism, as illustrated in Figure 9.

The design of the wing box structure for the HLFC demonstrator aimed to achieve a geometrically and structurally representative configuration, capable of accommodating the SPF/DB titanium perforated skin, the HLFC suction channels, the WIPS thermal circuit and the movable Krueger panel, while maintaining the aerodynamic quality necessary for laminarity validation tests. The 3D model of the assembly was used to define the boundary details and three-dimensional interfaces, while finite element analysis (FEA) was applied to critical junctions—including the front spar support, FLE interfaces and intermediate rib attachments—to justify material selection and sizing. These studies led to the freezing of the internal chamber layout, the definition of manifold routing and the specification of geometric features such as cut-outs and sealing interfaces with the kinematic mechanism.

In parallel, key design characteristics and tolerance chains were established to define the geometric datums used during manufacturing, ensuring a “no-fastener” aerodynamic surface in critical laminar flow regions. These design decisions were taken to minimize rework requirements and guarantee repeatable assembly within the ground test rig. Furthermore, the design process included the definition of rib-to-spar attachment solutions (blind-bolt joints for closed structures), titanium thermal shield placement in the WIPS zones and the integration of measurement and access points for instrumentation. All design aspects were structurally substantiated and verified through CAD and FEM analyses to ensure compliance with aerospace standards.

3. Manufacturing and Assembly Process

As part of the development of the Hybrid Laminar Flow Control (HLFC) leading-edge demonstrator, the manufacturing process constituted a critical phase in which the preliminary design concepts were transformed into viable technological solutions that satisfy aeronautical requirements for geometric precision, reduced mass and adequate structural integrity. At this stage, particular attention was devoted to the selection and implementation of advanced materials—namely, titanium alloys and carbon-fiber-reinforced polymer (CFRP) composites—as well as to the validation of cutting-edge fabrication technologies. Among these, Superplastic Forming combined with Diffusion Bonding (SPF/DB) was employed for the production of the perforated titanium suction skin, enabling the required micro-perforation quality and structural performance.

The manufacturing workflow was defined to ensure coherent integration of all primary subsystems, including the CFRP structure (wing box), the titanium suction skin mounted on LE internal structure made out of aluminium ribs, the Krueger panel and actuation mechanism and, last but not least, the internal anti-icing and suction titanium pipes. Each technological step was specified in detail to guarantee dimensional fidelity, thermal and electrical compatibility among materials and a streamlined approach to assembly and maintenance.

The subsequent sections describe the main steps of the demonstrator’s manufacturing process. These include the fabrication of the perforated titanium panel, entire leading edge structure-aluminium alloy ribs, the production of the CFRP structure, Krueger panel and associated kinematic arms and shaft, WIPS and suction systems and the final integration and assembly on the dedicated ground-test rig.

3.1. SPF/DB Titanium Perforated Skin Production Process

The fabrication of the suction panel for the Hybrid Laminar Flow Control (HLFC) system represented one of the most complex stages in the development of the demonstrator. The main objective was to obtain a double-curved titanium surface equipped with internal chambers capable of uniformly distributing suction along the entire leading-edge span (Figure 10).

This was achieved using the Superplastic Forming and Diffusion Bonding (SPF/DB) technology applied to titanium alloy sheets of Ti-6Al-4V (Figure 11).

The process began with the preparation of the titanium layers, which were chemically cleaned and locally coated with stop-off materials in the areas where diffusion bonding was to be prevented, in order to form the internal chambers. The superplastic forming was carried out in an argon atmosphere, at a temperature of approximately 900 °C and a pressure of 7 bar, followed by a vacuum diffusion bonding stage to achieve a complete metallurgical bond between the layers.

After bonding, the outer surface was laser-perforated, creating holes with diameters ranging from 80 to 120 µm, distributed at a 0.5 mm pitch, in order to ensure the required permeability for the HLFC suction system (Figure 12).

Post-processing included mechanical polishing to achieve a surface roughness of Ra ≤ 0.8 µm, a critical condition for maintaining laminar flow.

Dimensional inspections performed after bonding and machining indicated geometric deviations within ±0.15 mm, consistent with the precision requirements for aerodynamic surfaces. The internal integrity of the panel was verified through ultrasonic (C-scan) inspection and liquid penetrant testing, both confirming the absence of delamination and blockages within the suction channels.

In conclusion, the SPF/DB process proved to be reliable and reproducible, ensuring high structural quality. However, maintaining uniform perforation on curved surfaces required precise adjustments of the laser drilling parameters and accurate positioning within the forming fixture.

3.2. Manufacturing and Integration of the Suction and Wing Ice Protection Systems (WIPSs)

The manufacturing stage of the suction and ice protection systems represented a key component in the development of the HLFC demonstrator, aiming to reproduce the functional behavior of the hot-air and suction circuits for validating the Hybrid Laminar Flow Control concept for testing conditions. The design of these systems had already been finalized (including the suction chamber layout and manifold geometry), so the manufacturing activities focused on transforming the 3D models into high-precision metallic components, in compliance with applicable aerospace standards for pipes, fittings and thermal insulation elements.

The manufacturing processes began with the fabrication of the main pipes and manifolds made of titanium (Ti6Al4V), using high-precision CNC machining, followed by TIG welding in an inert argon atmosphere to ensure tightness. Each section was subsequently pressure-tested individually to verify the absence of leaks and the uniformity of the internal cross-section. In high-temperature areas, the pipes were equipped with local thermal insulation, designed to minimize heat loss and protect the leading-edge supporting structure (Figure 13).

Within the Wing Ice Protection System (WIPS) subassembly, transfer ducts and machined collectors mounted under the skin were manufactured to supply hot air to the critical regions of the titanium skin. These ducts were fitted with standardized couplings and connectors compliant with civil aircraft specifications.

For the validation of geometric integration, a “dummy valve”—a non-functional 3D-printed component—was produced to simulate the space allocated for the real part and to verify accessibility, segregation compliance and compatibility with cable routes and adjacent structures (Figure 14). This component allowed fine-tuning of assembly tolerances and served as a reference element during the final assembly process.

The final assembly of the WIPS and HLFC systems was carried out in parallel with the integration of the metallic leading-edge structure. The pipes were positioned in the jig according to the integration drawings, fixed using aluminum clamps and elastomeric seals and all connections were dimensionally and functionally verified using compressed air tests at pressures between 0.5 and 1.5 bar. At the same time, the air collection and distribution components of the HLFC system—also made of titanium—were installed to ensure uniform flow through the perforated skin chambers.

In the final stage, the complete system was equipped with pressure and flow sensors for initial calibration, and its compatibility with the air supply installation (previously described) was verified. The result of the manufacturing process was a compact, easy-to-maintain assembly, fully integrated into the leading-edge structure and fully compliant with segregation requirements between the pneumatic, thermal and electrical circuits. The quality of workmanship was confirmed through leak tests and non-destructive inspections (NDTs) performed at INCAS, which demonstrated compliance with design requirements and dimensional stability throughout assembly.

3.3. Leading Edge Structure Manufacturing and Assembly

The internal structure of the leading edge consists of the primary ribs, the Bird Impact Protection (BIP) system, the forward and lower attachment spars, as well as titanium thermal shields designed to isolate the regions exposed to the hot-air flow of the de-icing system (Figure 15).

The ribs, stoppers and spacers were machined from 7050-T7651 aluminum alloy, selected for its combination of mechanical strength, dimensional stability and low density. Components subjected to impact loading were manufactured from 7475-T761 aluminum alloy with a nominal thickness of 4.5 mm. The mounting brackets were formed through a deep-drawing process followed by CNC finishing on five-axis machining centers and subsequently hard-anodized to ensure corrosion protection.

For the assembly of the structure, liquid shims were applied between the ribs and the CFRP components (spar and skin) to compensate for thickness variations and to ensure uniform distribution of contact loads. The ribs were attached to the spar using blind bolts, a solution essential for closed-section structures. Following the integration of the Bird Impact Protection System, the titanium sheet thermal shields—machined to the CAD contour and cold-formed—were installed, followed by the placement of silicone sealing gaskets along the interface with the Krueger panel. The entire assembly was dimensionally inspected in the integration jig with an accuracy of ±0.1 mm, in accordance with the design requirements.

3.4. Krueger Flap and Kinematics Fabrication and Integration + Actuation Shaft

The Krueger panel is manufactured from carbon-fiber-reinforced polymer (CFRP) in a sandwich configuration incorporating a structural foam core, providing high bending stiffness at minimal weight. This architecture was selected for its favorable strength-to-weight ratio, superior fatigue resistance and dimensional stability under moderate thermal gradients. During the detailed design phase, the external geometry of the panel was refined to ensure aerodynamic continuity with the fixed leading-edge contour and to minimize gaps in the retracted position that could disturb the local flow.

The actuation system, fabricated entirely from high-strength 15-5 PH stainless steel, employs a kinematic arrangement that produces a controlled rotational deployment. The mechanism consists of a primary actuator lever, a drive-link shaft and an articulated guiding linkage that together allow smooth motion throughout the deployment and retraction cycle without interfering with the surrounding wing structure (Figure 16).

The kinematics were further optimized to guarantee a positional alignment within 0.2 mm in the stowed configuration—a critical requirement to ensure proper sealing and to preserve laminar-flow continuity across the upper surface of the leading edge.

The actuation shaft is basically a segmented shaft made from standard parts (all the interfaces are) and machined ones—steel was used to sustain the high stress values that will occur during the icing test and bird strike test. The gearboxes in our case are only designed to fulfill the space allocation model; in reality, these parts house the bearings for the actuation chain.

The manufacturing processes of the Krueger mechanism followed the same precision standards as those applied to the internal structure. All major components were manufactured from high-strength steel [28] (EN 34CrNiMo6), machined on 5-axis CNC centers, while the contact surfaces were finished by grinding and lapping, achieving a concentricity below 0.02 mm and a surface roughness of Ra ≤ 0.8 µm.

After the hardening heat treatment (38–42 HRC), the components were cadmium-plated and passivated, while the aluminum parts were hard-anodized. The bearing–lever–shaft subassemblies were manually assembled using pneumatically controlled presses with force monitoring, and the self-lubricating bushings were installed by hot pressing to ensure dimensional stability.

3.5. Wing Box Manufacturing

Wing Box subscales demonstrators have been prepared and implemented in order to de-risk the concept for high-quality aero surface (no fastener) composite wing structures (Figure 17).

The manufacturing process of the wing box began with the definition of tooling and assembly jigs based on the final CAD model, to ensure the geometric referencing (“datums”) required for the subsequent assembly of subassemblies. The tooling was designed to maintain the positioning of the front spar and the panel edges within the tolerance limits imposed by aerodynamic analyses and by the “no-fastener surface” requirement on the external surfaces.

The fabrication of the CFRP components followed conventional manual and automated lay-up processes on molds conforming to the final contour. Prepreg materials were applied to mandrels and forming meshes, followed by autoclave curing cycles in accordance with the approved thermal parameters. After demolding, trimming and mechanical finishing operations were performed to obtain the final edges and assembly interfaces. Layer thickness and fiber/resin orientation were verified post-cure through 3D measurements and metrological inspection procedures.

The internal metallic components (ribs, stoppers, spacers) were CNC-machined from aluminum alloys (Al 7050-T7651 for the ribs and Al 7475-T761 for bird-impact components), including milling, drilling and grinding operations performed to tight tolerances. After machining, these parts underwent surface treatments (anodizing or passivation) and, where required, bending/forming operations followed by final finishing. Critical features (punch holes, attachment flanges) were verified through CAD-to-part comparison prior to release for assembly.

3.6. Integration of the Components and Final Assembly of the HLFC Wing Demonstrator

The final assembly of the Hybrid Laminar Flow Control (HLFC) demonstrator was carried out at INCAS using a dedicated high-tolerance assembly jig specifically designed to ensure accurate three-dimensional positioning of all major subassemblies. These included the CFRP wing-box structure, the aluminum leading edge ribs, the SPF/DB-formed titanium suction panel and the CFRP Krueger panel and stainless steel kinematic system. The jig incorporated micrometric clamping systems and adjustable transverse supports (0.1 mm increment), while its rigid steel frame was calibrated prior to use to eliminate elastic deformations during component handling (Figure 18).

The assembly process followed a staged workflow. It began with the positioning and fixation of the CFRP wing box, which provided the geometric reference for the entire demonstrator. Leading-edge metallic ribs and the bird-impact-protection (BIP) sub-spar were subsequently installed on this baseline structure, followed by the routing and placement of the HLFC and WIPS system ducts. All pneumatic and thermal interfaces were initially sealed using temporary fittings to facilitate further positional adjustment and preliminary functional checks. The ducts and manifolds were secured with aluminum brackets and silicone insulating inserts to prevent vibration transfer into the structure.

In parallel, the Krueger kinematic mechanism was integrated and bench-tested in its extreme deployed and retracted positions, demonstrating smooth operation without interference between moving elements and support structure.

Following completion of the structural assembly, the pressure, flow and temperature sensors for the HLFC and WIPS systems were installed. All electrical connections were routed along shielded, electromagnetically protected pathways to prevent interference during functional testing.

After completion of the internal pipe system integration, the SPF/DB-manufactured perforated titanium panel was installed on the leading edge. This step represented one of the most critical phases, as the external surface had to remain flush and uninterrupted (“no-fastener surface”) to preserve laminar-flow conditions. The panel was joined using blind fasteners and silicone gaskets rated for temperatures up to 200 °C. Dimensional verification performed using a CREAFORM measurement system confirmed installation tolerances within ±0.2 mm.

The final inspection results confirmed full compatibility among all integrated subsystems and compliance with the design specifications. Metrological measurements indicated a total mass of 311 kg and a center-of-gravity location of x = 42.455 mm, y = −22.850 mm and z = 839 mm, in agreement with the digital model requirements. Assembly using the dedicated jig demonstrated not only the feasibility of multi-material integration (CFRP, titanium, aluminum and steel) but also the dimensional stability necessary for functional validation of the HLFC system under laboratory test conditions.

In the final phase, the fully assembled demonstrator was connected to the compressed-air supply installation and to the 35 kW electric heating system, enabling calibration procedures and leak-tightness verification (Figure 19).

4. Instrumentation and Data Acquisition

The instrumentation of the Ground Based Demonstrator (GBD) was defined to support the monitoring and validation of the Krueger flap kinematics, the Wing Ice Protection System (WIPS) and the HLFC suction system during ground and wind tunnel testing. Given the Research and Technology nature of the demonstrator, the instrumentation strategy focused on robustness, flexibility and reliable data acquisition rather than on full aerodynamic performance assessment.

Instrumentation included pressure sensors, thermocouples, flow meters and position sensors installed at key locations within the WIPS ducting, suction system collectors and trunk ducts and the Krueger kinematic chain. These sensors enabled the measurement of pressure, temperature, mass flow and actuator position under steady-state and transient operating conditions,

Table 1. Instrumentation overview.

Component	Specific Measurement Objective [Unit]	Sensor Type	Sensor Spec	Sensor Range	Sensor Tolerance	Sensor Position
Nose connector	Sucked air pressure [Pa]	Pressure taps	NSA 93959859	0 to 1.7 bar	0.08% Full Scale	Middle of the pipe
Top connector	Sucked air pressure [Pa]	Pressure taps	NSA 93959859	0 to 1.7 bar	0.08% Full Scale	Middle of the pipe
Trunk duct	Sucked air pressure [Pa]	Pressure taps	NSA 93959859	0 to 1.7 bar	0.08% Full Scale	Outboard zone
Trunk duct	Sucked air pressure [Pa]	Pressure taps	NSA 93959859	0 to 1.7 bar	0.08% Full Scale	Middle span
Trunk duct	Sucked air pressure [Pa]	Pressure taps	NSA 93959859	0 to 1.7 bar	0.08% Full Scale	Inboard zone
Nose connector	Sucked air temperature [°C]	Thermocouple OMEGA TT-K-30-SLE (TC type K)	FineWire_DuplusI Tc.pdf	0–150 °C	0.5 °C	Middle of the pipe
Trunk duct	Sucked air temperature [°C]	Thermocouple OMEGA TT-K-30-SLE (TC type K)	FineWire_DuplusI Tc.pdf	0–150 °C	0.5 °C	Middle span
Trunk duct	Sucked air temperature [°C]	Thermocouple OMEGA TT-K-30-SLE (TC type K)	FineWire_DuplusI Tc.pdf	0–150 °C	0.5 °C	Inboard zone
Inboard rib	Air temperature [°C]	Thermocouple OMEGA TT-K-30-SLE (TC type K)	FineWire_DuplusI Tc.pdf	0–150 °C	0.5 °C	Inboard zone
Middle rib	Air temperature [°C]	Thermocouple OMEGA TT-K-30-SLE (TC type K)	FineWire_DuplusI Tc.pdf	0–150 °C	0.5 °C	Middle span
Outboard rib	Air temperature [°C]	Thermocouple OMEGA TT-K-30-SLE (TC type K)	FineWire_DuplusI Tc.pdf	0–150 °C	0.5 °C	Outboard zone
Intermediate rib 1 (inboard)	Structure temperature [°C]	Thermocouple OMEGA SA1XL-K	OMEGA SA1XL-K	0–260 °C	0.5 °C	Close to rib attachments
Intermediate rib 3 (inboard)	Structure temperature [°C]	Thermocouple OMEGA SA1XL-K	OMEGA SA1XL-K	0–260 °C	0.5 °C	Close to rib attachments

and Figure 20.

A centralized data acquisition system (DAQ) was implemented to ensure synchronized recording of all measurement channels. The DAQ architecture allowed real-time monitoring and logging of sensor data during deployment cycles, pressure ramps and thermal tests. Sampling rates and sensor accuracy were selected to capture both quasi-steady conditions and dynamic effects associated with actuation and system response.

Figure 20 illustrates the instrumentation layout of the skin, highlighting the arrangement of thermocouples and pressure taps used for temperature and pressure measurements during testing. The OMEGA thermocouples achieve response times of less than 0.15 s. These thermocouples are constructed using a Polyimide/fiberglass-junction insulation along with a fiberglass-insulated lead wire and can be used as a self-adhering thermocouple for temperatures up to 260 °C (500 °F), or they can be cemented in place for use at temperatures up to 315 °C (600 °F).

Specs: Dimensions:

Patch Length: 25.4 mm (1.0″);

Patch Width: 9.5 mm (0.375″);

Strip Length: 25.4 mm (1.0″) with 12.7 mm (0.5″) bare wire.

5. Ground Testing (Vacuum and Hot Air) Results

Pneumatic Test Bench (Heated Air and Vacuum)

For the purpose of ground-based functional testing of the Hybrid Laminar Flow Control (HLFC) and Wing Ice Protection System (WIPS), a complex air supply installation was designed to reproduce the pressure, temperature and flow rate conditions characteristic of real flight regimes. The installation design was developed by INCAS, aiming for seamless integration within a laboratory environment and full compatibility with the HLFC demonstrator testing infrastructure (Figure 21).

The installation was conceived with modular architecture to ensure adaptability to multiple experimental scenarios. From a functional point of view, it includes a compression system, an air conditioning unit, an active heating section and a real-time control and monitoring system. The two compressors can operate either independently or in parallel, reaching pressures up to 18 bar and maximum flow rates of approximately 620 kg/h. The compressed air is subsequently passed through a purification filter and then directed to the conditioning unit, where temperature and pressure are adjusted to match the nominal testing conditions.

The central element of the installation is a 35 kW electric heater, capable of increasing the air temperature from −30 °C up to +200 °C, while maintaining thermal stability throughout the entire operating range. The system was dimensioned to allow testing of the HLFC and WIPS components under variable regimes, ranging from cold, low-pressure flow conditions to intense thermal regimes simulating in-flight de-icing scenarios (Figure 22).

From an architectural standpoint, the pneumatic circuit consists of stainless-steel and titanium pipes and fittings, locally insulated in high-temperature areas. The configuration of the pipelines was analyzed in detail within a 3D CAD environment to minimize pressure losses and optimize flow uniformity at the outlet of the heating system. Additionally, to prevent any risk of overpressure or overheating, the installation was equipped with redundant safety systems, including a passive thermal protection mechanism that shuts down the power supply when the temperature exceeds 90 °C and an electronic control system that interrupts operation in the absence of airflow or under excessive flow rate conditions.

The WIPS test was performed at the following pressures: 1.6 bar, 1.4 bar, 1.2 bar, 1 bar, 0.8 bar, 0.6 bar, 0.4 bar and 0.2 bar. The results are presented in

Table 2. WIPS results.

Point	1	2	3	4	5	6	7	8	9	10	11	12	13	14	15	16	17	18	19	20
WIPS test at pressure 1.6 bar, measured flow [kg/h] = 584.797
Temperature [°C]	30.94	33.253	31.175	33.511	33.161	32.361	32.635	31.47	33.124	31.527	30.538	29.253	32.327	-	-	30.806	30.461	20.482	-	-
Pressure [bar]	1.302	1.310	1.295	1.506	1.559	1.580	1.487	1.474	0.017	0.001	0.000	0.001	0.001	1.587	1.586	1.585	0.000	1.311	1.600	1.533
WIPS test at pressure 1.4 bar, measured flow [kg/h] = 554.545
Temperature [°C]	27.233	29.591	28.501	29.754	29.632	29.255	28.876	28.19	29.443	28.626	28.359	27.79	29.095	-	-	28.312	28.128	20.45	-	-
Pressure [bar]	1.109	1.116	1.102	1.288	1.338	1.359	1.275	1.264	0.014	0.000	0.000	0.000	0.000	1.365	1.364	1.363	0.000	1.117	1.377	1.314
WIPS test at pressure 1.2 bar, measured flow [kg/h] = 558.994
Temperature [°C]	25.579	27.982	27.244	28.079	28.064	27.832	27.191	26.705	27.831	27.325	27.267	26.977	27.65	-	-	27.116	27.019	20.33	-	-
Pressure [bar]	1.044	1.051	1.038	1.215	1.263	1.284	1.205	1.193	0.012	0.000	0.000	0.000	0.000	1.290	1.289	1.288	0.000	1.052	1.302	1.240
WIPS test at pressure 1 bar, measured flow [kg/h] = 491.554
Temperature [°C]	20.839	23.091	23.179	23.128	23.246	23.344	22.305	22.324	22.945	23.018	23.466	23.867	23.129	-	-	23.188	23.29	19.951	-	-
Pressure [bar]	0.787	0.791	0.781	0.919	0.963	0.983	0.918	0.908	0.009	0.000	0.000	0.000	0.000	0.988	0.987	0.987	0.000	0.792	0.998	0.943
WIPS test at pressure 0.8 bar, measured flow [kg/h] = 483.124
Temperature [°C]	19.929	22.026	22.215	22.072	22.186	22.32	21.313	21.362	21.88	22.064	22.543	23.078	22.102	-	-	22.244	22.354	19.756	-	-
Pressure [bar]	0.692	0.697	0.687	0.810	0.851	0.871	0.812	0.802	0.007	0.000	0.000	0.000	0.000	0.875	0.875	0.874	0.000	0.697	0.884	0.832
WIPS test at pressure 0.6 bar, measured flow [kg/h] = 443.915
Temperature [°C]	19.116	20.929	21.286	20.937	21.106	21.326	20.249	20.354	20.795	21.091	21.589	22.282	21.094	-	-	21.319	21.467	19.558	-	-
Pressure [bar]	0.468	0.471	0.464	0.546	0.581	0.600	0.557	0.549	0.005	0.000	0.000	0.000	0.000	0.603	0.603	0.602	0.000	0.472	0.610	0.566
WIPS test at pressure 0.4 bar, measured flow [kg/h] =376.145
Temperature [°C]	19.056	20.341	20.613	20.401	20.426	20.585	19.922	19.855	20.234	20.496	20.924	21.578	20.499	-	-	20.685	20.855	19.528	-	-
Pressure [bar]	0.331	0.334	0.328	0.383	0.413	0.429	0.398	0.391	0.003	0.000	0.000	0.000	0.000	0.432	0.432	0.431	0.000	0.334	0.437	0.400
WIPS test at pressure 0.2 bar, measured flow [kg/h] = 236.506
Temperature [°C]	18.433	19.172	19.426	19.146	19.183	19.367	18.895	18.804	19.047	19.385	19.722	20.357	19.313	-	-	19.538	19.717	19.09	-	-
Pressure [bar]	0.174	0.175	0.172	0.192	0.214	0.227	0.211	0.207	0.001	0.000	0.000	0.000	0.000	0.229	0.229	0.228	0.000	0.175	0.232	0.204

.

The measurements vacuum test was performed in three configurations at the following absolute pressures:

First Configuration (orifice place—small diam.): 0.9, 0.8, 0.7, 0.6 and 0.5 bar;

Table 3. Measurements vacuum test in first configuration.

Point	1	2	3	4	5	6	7	8	9	10	11	12	13	14	15	16	17	18	19	20
pressure 0.9 bar, measured flow [kg/h] = 215.349
Temperature [°C]	17.325	17.226	17.255	17.285	17.253	17.243	17.121	17.012	17.129	17.152	17.166	17.165	16.976	-	-	17.213	17.166	17.212	-	-
Pressure [bar]	0.939	0.937	0.937	0.955	0.945	0.936	0.937	0.937	0.898	0.897	0.913	0.896	0.896	0.937	0.937	0.936	0.895	0.898	0.937	0.950
pressure 0.8 bar, measured flow [kg/h] = 213.468
Temperature [°C]	17.33	17.233	17.215	17.247	17.221	17.209	17.097	16.999	17.053	17.103	17.123	17.166	16.844	-	-	17.153	17.116	17.059	-	-
Pressure [bar]	0.898	0.895	0.896	0.921	0.907	0.894	0.895	0.896	0.815	0.813	0.847	0.810	0.810	0.895	0.895	0.895	0.809	0.821	0.895	0.913
pressure 0.7 bar, measured flow [kg/h] = 312.120
Temperature [°C]	17.25	17.161	17.162	17.19	17.172	17.159	17.013	16.983	16.999	17.077	17.199	17.086	16.705	-	-	17.139	17.087	16.547	-	-
Pressure [bar]	0.862	0.858	0.859	0.888	0.871	0.855	0.857	0.859	0.713	0.709	0.768	0.703	0.703	0.856	0.856	0.856	0.702	0.750	0.857	0.878
pressure 0.6 bar, measured flow [kg/h] = 326.204
Temperature [°C]	17.182	17.028	17.09	17.093	17.064	17.04	16.991	16.868	16.912	16.983	17.022	16.988	16.367	-	-	17.014	16.927	14.787	-	-
Pressure [bar]	0.847	0.844	0.846	0.875	0.855	0.839	0.842	0.844	0.613	0.607	0.695	0.598	0.597	0.840	0.840	0.840	0.596	0.729	0.842	0.863
pressure 0.5 bar, measured flow [kg/h] = 336.286
Temperature [°C]	17.155	17.028	17.074	17.072	17.068	17.029	16.977	16.878	16.906	16.988	17.022	16.988	16.233	-	-	17.022	16.904	13.922	-	-
Pressure [bar]	0.842	0.839	0.841	0.870	0.849	0.832	0.836	0.838	0.560	0.552	0.656	0.541	0.539	0.834	0.834	0.833	0.538	0.720	0.836	0.857

.

Second Configuration (WIPS duct blocked—full plate)

Table 4. Measurements vacuum test in second configuration.

Point	1	2	3		4	5	6	7	8	9	10	11	12	13	14	15	16	17	18	19	20
pressure 0.9 bar, measured flow [kg/h] = 138.766
Temperature [°C]	16.932	16.832	17.336		16.917	17.017	17.034	16.744	16.763	16.806	16.904	17.009	16.744	16.827	-	-	17.11	16.941	16.997	-	-
Pressure [bar]	0.983	0.975	0.974		0.995	0.995	0.995	0.985	0.982	0.906	0.904	0.918	0.903	0.902	0.993	0.994	0.994	0.902	0.903	0.989	0.999
pressure 0.8 bar, measured flow [kg/h] = 235.096
Temperature [°C]	16.921	16.744	17.101		16.9	16.895	16.878	16.728	16.762	16.797	16.848	16.764	16.757	16.667	-	-	17.011	16.858	16.828	-	-
Pressure [bar]	0.971	0.956	0.954		0.991	0.991	0.992	0.975	0.970	0.813	0.809	0.843	0.806	0.806	0.988	0.989	0.990	0.806	0.817	0.981	0.998
pressure 0.7 bar, measured flow [kg/h] = 302.025
Temperature [°C]	16.859	16.639	17.336		16.838	16.845	16.845	16.686	17.019	16.781	16.798	16.759	16.724	16.605	-	-	16.92	16.808	15.781	-	-
Pressure [bar]	0.966	0.949	0.946		0.989	0.989	0.991	0.971	0.965	0.706	0.699	0.760	0.694	0.693	0.986	0.988	0.989	0.692	0.781	0.978	0.998
pressure 0.6 bar, measured flow [kg/h] = 334.644
Temperature [°C]	16.823	16.622	17.044		16.794	16.754	16.734	16.648	16.61	16.709	16.692	16.603	16.678	16.488	-	-	16.836	16.671	13.674	-	-
Pressure [bar]	0.966	0.950	0.947		0.989	0.990	0.991	0.971	1.034	0.613	0.603	0.690	0.594	0.593	0.936	0.988	0.989	0.592	0.785	0.979	0.998
pressure 0.5 bar, measured flow [kg/h] = 340.975
Temperature [°C]	16.826	16.646	17.021		16.812	16.772	16.755	16.646	16.657	16.714	16.698	16.619	16.682	16.502	-	-	16.932	16.625	12.991	-	-
Pressure [bar]	0.998	0.997	0.996		0.999	0.999	0.999	0.998	0.998	0.968	0.967	0.975	0.967	0.967	0.999	0.999	0.999	0.967	0.986	0.999	1.000

.

Third Configuration (micro-perforated holes sealed—adhesive tape on stripes).

Table 5. Measurements vacuum test in third configuration.

Point	1	2	3	4	5	6	7	8	9	10	11	12	13	14	15	16	17	18	19	20
pressure 0.9 bar, measured flow [kg/h] = 158.379
Temperature [°C]	17.672	17.661	17.639	17.569	17.535	17.518	17.442	17.356	17.362	17.295	17.301	17.291	17.311	-	-	17.270	17.399	17.463	-	-
Pressure [bar]	0.928	0.926	0.925	0.948	0.936	0.926	0.927	0.929	0.908	0.908	0.923	0.906	0.907	0.927	0.927	0.927	0.905	0.907	0.927	0.942
pressure 0.8 bar, measured flow [kg/h] = 229.924
Temperature [°C]	17.745	17.705	17.754	17.645	17.560	17.541	17.473	17.373	17.303	17.282	17.299	17.278	17.316	-	-	17.315	17.488	17.577	-	-
Pressure [bar]	0.860	0.856	0.855	0.891	0.873	0.858	0.859	0.864	0.810	0.808	0.846	0.805	0.805	0.858	0.859	0.858	0.804	0.811	0.858	0.881
pressure 0.7 bar, measured flow [kg/h] = 265.871
Temperature [°C]	17.626	17.621	17.693	17.612	17.538	17.513	17.449	17.323	17.225	17.239	17.297	17.274	17.160	-	-	17.311	17.446	17.175	-	-
Pressure [bar]	0.802	0.797	0.797	0.842	0.819	0.799	0.802	0.809	0.714	0.705	0.773	0.705	0.705	0.800	0.801	0.800	0.703	0.728	0.801	0.829
pressure 0.6 bar, measured flow [kg/h] = 276.435
Temperature [°C]	17.644	17.683	17.783	17.649	17.550	17.530	17.506	17.397	17.229	17.252	17.302	17.293	17.133	-	-	17.339	17.489	16.648	-	-
Pressure [bar]	0.762	0.756	0.756	0.808	0.781	0.758	0.762	0.770	0.621	0.617	0.705	0.607	0.607	0.759	0.760	0.759	0.605	0.673	0.760	0.791
pressure 0.5 bar, measured flow [kg/h] = 293.808
Temperature [°C]	17.653	17.617	17.764	17.704	17.633	17.595	17.541	17.300	17.268	17.313	17.389	17.377	16.974	-	-	17.428	17.400	15.224	-	-
Pressure [bar]	0.731	0.725	0.725	0.781	0.751	0.726	0.731	0.741	0.516	0.509	0.632	0.495	0.492	0.728	0.728	0.727	0.491	0.634	0.729	0.744

. Cells marked with “–” indicate locations where thermocouples were not installed, and, therefore, no measurements were taken.

The pressure range of 1.6–0.2 bar was selected to cover the operational envelope of the Wing Ice Protection System (WIPS) and the HLFC suction system for ground test conditions, and this range was used also for future calibration of the virtual tool that was used for this study.

Table 2 summarizes the experimental results obtained during the ground tests of the Wing Ice Protection System (WIPS) under hot-air operating conditions, for supply pressures ranging from 1.6 bar down to 0.2 bar. The measurements include local temperatures and pressures recorded at multiple locations along the leading-edge structure, as well as the corresponding mass flow rates.

The results indicate a high level of thermal uniformity along the protected region. For the highest operating pressure (1.6 bar), the measured temperatures are predominantly within a narrow band of around 31–33 °C, with localized lower values near the extremities, attributable to boundary effects and heat losses. As the supply pressure decreases, a gradual reduction in temperature is observed, following the expected thermodynamic behavior of the system.

Pressure measurements confirm a stable and well-distributed pressure field within the WIPS ducting, with negligible pressure drops between adjacent measurement points. This behavior demonstrates the effectiveness of the internal duct layout and validates the design assumptions regarding hot-air distribution.

Overall, the data in Table 2 confirm that the WIPS is capable of delivering consistent thermal performance over a wide operating envelope, without inducing excessive pressure losses or temperature gradients that could compromise structural integrity or laminar-flow compatibility.

Table 3 presents the results of the vacuum tests performed in the first configuration, corresponding to the reference suction layout with the standard orifice geometry. Measurements were conducted at absolute pressures between 0.9 bar and 0.5 bar, covering the representative operational range of the HLFC suction system.

The recorded temperatures remain nearly constant, with variations typically below ±0.2 °C, indicating that suction-induced cooling effects are minimal and do not introduce thermal disturbances. This behavior is particularly important for HLFC applications, where temperature gradients can affect boundary-layer stability.

Pressure measurements show a consistent pressure decay along the suction path, with smooth gradients and no evidence of flow separation or localized choking. The measured mass flow rates increase as the absolute pressure decreases, confirming the expected inverse relationship between suction pressure and flow rate.

These results validate the baseline suction configuration, demonstrating stable operation, predictable flow behavior and compatibility with the HLFC design requirements.

Table 4 reports the vacuum test results obtained in the second configuration, in which the WIPS duct was blocked, isolating the suction system to evaluate its intrinsic behavior without interaction with the hot-air circuit.

Compared to Configuration 1, the measured mass flow rates are generally lower at higher pressures, reflecting the modified boundary conditions imposed by the blocked duct. However, as the pressure decreases toward 0.6–0.5 bar, the flow rates converge toward values comparable to those observed in the reference configuration.

Temperature measurements again exhibit very limited variation, confirming that the suction process remains thermally benign. The pressure data show a more uniform distribution across several measurement points, indicating reduced interaction effects and validating the internal suction chamber design.

This configuration demonstrates that the HLFC suction system can operate independently of the WIPS circuit, an important finding for system robustness and failure-tolerance considerations.

Table 5 summarizes the results of the vacuum tests conducted in the third configuration, where the micro-perforated holes were sealed. This setup was intended to characterize parasitic losses and baseline system behavior in the absence of active suction through the skin.

As expected, the measured mass flow rates are significantly reduced compared to Configurations 1 and 2, confirming that the majority of the airflow in previous tests is effectively driven through the perforated titanium skin. The pressure levels measured along the circuit show a more homogeneous distribution, with reduced gradients, reflecting the absence of distributed suction sources.

Temperature values remain stable and comparable to the other configurations, further supporting the conclusion that thermal effects are decoupled from suction behavior under the tested conditions.

The results of Table 5 provide a reference baseline, allowing clear discrimination between active suction effects and system-intrinsic losses. This comparison reinforces the validity of the suction concept and supports the interpretation of the results obtained in the previous configurations.

Taken together, the results presented in Table 2, Table 3, Table 4 and Table 5 demonstrate the stable, predictable and robust operation of both the Wing Ice Protection System and the HLFC suction network over a wide range of operating conditions. The experimental data confirm the effectiveness of the integrated design, validate the underlying numerical assumptions and provide a solid experimental foundation for subsequent icing wind tunnel and aerodynamic performance tests.

6. Wind Tunnel Testing (Icing/Anti-Icing) Results

6.1. Facility and Setup

The aerodynamic and icing experiments were performed in the Icing Wind Tunnel (IWT) operated by CIRA, as illustrated in Figure 23. The IWT is a closed-circuit, Göttingen-type wind tunnel, equipped with a refrigeration system and the capability to operate under both reduced and elevated pressure, allowing the simulation of icing conditions corresponding to specific flight altitudes. Icing clouds are generated by means of spray bars, which enable controlled injection of water droplets with prescribed size characteristics.

Among the available test sections, the facility segment with a cross-section of 2.35 m × 2.25 m and a length of 7 m was selected for the present campaign. Under these conditions, the wind tunnel can reach flow velocities up to Mach 0.41 while operating at static temperatures as low as −32 °C.

A more detailed description is made by Vecchione et al. [29], where a dedicated presentation of this unique testing facility is presented, highlighting the main characteristics and capabilities for icing conditions and test item dimensions.

For the wind tunnel campaign, supplementary instrumentation was implemented. The Krueger flap was instrumented with fourteen static pressure ports to measure local surface pressures and to provide validation of the aerodynamic design. Additionally, a pressure belt installed on the wing was used to adjust the trim flap, ensuring the target pressure distribution and compensating for the reduced chord length of the dummy wing box. Key wind tunnel parameters—including flow velocity, temperature, humidity and pressure—were continuously monitored and recorded by the facility’s data acquisition system. The Krueger flap was also monitored with respect to actuation torque and angular displacement.

6.2. High-Lift Icing Test

In the current configuration, the Krueger flap was not equipped with a dedicated ice protection system. One of the primary objectives of the test campaign was to assess whether the absence of such a system would have a detrimental impact on the overall design. Introducing an ice protection system on the Krueger flap would significantly increase installation complexity within the already congested leading-edge region and would likely require a shift from hot-air anti-icing to an electrical de-icing solution. Supplying sufficient hot air to the thin, deployed Krueger panel—together with the associated ducting—was considered impractical. Therefore, demonstrating that an additional ice protection system was unnecessary would represent a substantial simplification of the integration concept.

Icing accretion tests were conducted with the Krueger flap deployed under representative approach conditions, corresponding to an angle of attack of α = 6° and a free-stream Mach number of M∞ = 0.23. The icing environment was defined by a static temperature of −23 °C and two liquid water content levels (LWC = 0.3 g/m³ and 0.4 g/m³), associated with median volumetric droplet diameters (MVDs) of 50 μm and 20 μm, respectively.

The Ground-Based Demonstrator was subjected to these icing conditions for a duration of 13 min, representative of typical take-off and landing phases with the high-lift system deployed. In the most severe case, ice accretion on the Krueger flap reached thicknesses of up to 10 mm, primarily in the form of rough ice, with no evidence of horn-ice formation. Crucially, no ice deposits or signs of water impingement were observed on the upper surface of the main wing, confirming the effective shielding role of the Krueger flap in protecting the laminar surface. Limited ice residues were detected within the lower cavity region.

Post-test actuation of the Krueger flap demonstrated that neither the accumulated ice on the flap nor the residual ice in the cavity hindered its motion within the predefined torque limits. These results [27] indicate that the Krueger flap effectively prevents ice and contamination from reaching the main wing laminar surface, while remaining fully functional even after exposure to representative icing conditions over the selected duration.

6.3. GBD-Icing Test

Following the icing tests described in the previous section, an additional assessment was performed to evaluate the behavior of the Krueger flap (retracted—clean configuration) and the leading-edge region after exposure to representative icing conditions. The objective of this analysis was to assess the extent of ice accretion on the high-lift device and to verify whether the accumulated ice affects either the mechanical operability of the system or the shielding capability of the laminar leading-edge surface.

Figure 24 presents representative images of the Krueger flap (retracted) and the leading-edge region immediately after the icing exposure in the Icing Wind Tunnel. The figure indicates the ice formation thickness, and it was measured using a hot plate template.

The visual inspection indicates limited ice accretion on the leading edge area, mainly in the form of thin rough ice layers, with no tendency for horn-ice formation. In Figure 25, no ice accretion, water impingement marks or residual contamination were observed on the main wing upper surface, confirming the effective shielding capability of the deployed Krueger flap. Minor ice residuals were detected only within the lower cavity and did not interfere with the kinematic mechanism or sealing interfaces.

Post-test actuation of the Krueger flap confirmed that the observed ice accretion did not prevent deployment or retraction of the mechanism and that the required actuation torque remained within the assumed limits. From both an aerodynamic and mechanical perspective, the ice accretion was therefore assessed as non-critical.

These results demonstrate that the Krueger flap effectively protects the laminar leading-edge surface against icing and contamination and remains fully operational after representative icing exposure. Consequently, the findings support the conclusion that a dedicated ice protection system for the Krueger flap itself is not required under the tested conditions.

7. Conclusions

This study presented the development and experimental validation of a full-scale Hybrid Laminar Flow Control (HLFC) wing leading-edge demonstrator integrating suction-based laminar-flow technology, a hot-air Wing Ice Protection System (WIPS) and a folding bull-nose Krueger high-lift device. The Ground-Based Demonstrator was designed to realistically reproduce the geometric, structural and functional constraints of a transport aircraft’s outer-wing leading edge while allowing experimental flexibility for research and validation activities.

The manufacturing and integration campaign successfully demonstrated the feasibility of combining advanced materials and processes—such as CFRP structures, SPF/DB-manufactured perforated titanium skins and precision-machined metallic components—into a compact and robust leading-edge assembly. Ground-based functional testing confirmed the stable and uniform thermal performance of the WIPS over a wide range of operating pressures, as well as the predictable and repeatable behavior of the HLFC suction system under various configurations.

Icing wind tunnel tests conducted at the CIRA Icing Wind Tunnel Facility further demonstrated the effectiveness of the Krueger flap as a shielding device for the laminar leading-edge surface. Ice accretion on the deployed Krueger flap was limited and non-critical, with no horn-ice formation observed and no contamination detected on the main wing upper surface. Post-test actuation verified that the accumulated ice did not impair the mechanical operability of the high-lift system, supporting the conclusion that a dedicated ice protection system for the Krueger flap itself is not required under the tested conditions.

Overall, the experimental results confirm that the proposed integrated HLFC leading-edge architecture provides a viable and robust solution for combining laminar-flow control, ice protection and high-lift functionality within the limited space of a transport aircraft wing leading edge.

Future work will focus on a detailed analysis of the aerodynamic and icing wind tunnel test results. This will include a comprehensive evaluation of pressure distributions, flow-field characteristics and laminar-flow behavior under a wider range of angles of attack, Mach numbers and icing conditions. Particular attention will be given to correlating experimental measurements with numerical simulations in order to further assess aerodynamic performance, transition behavior and the robustness of the HLFC concept under realistic operational scenarios. These results will be presented in a dedicated follow-up publication.