Abstract: This chapter explores aluminium alloys' vital role in next-gen defense systems, focusing on their evolution, processing, and aerospace/military applications. It traces their development from duralumin to advanced Al-Li and nanostructured variants, emphasizing strength-to-weight efficiency and corrosion resistance. Key processing innovations like thermomechanical treatments, additive manufacturing, and friction stir welding enable lightweight, high-performance components, such as AlSi10Mg brackets in the F-35 and GLARE® fuselage panels in the A380. Case studies highlight alloys like AA 7085 and Al-Li 2195 in military aircraft and spacecraft, enhancing fuel efficiency and durability. Challenges like galvanic corrosion, scandium scarcity, and energy-intensive production are analyzed, alongside solutions like AI-driven alloy design and closed-loop recycling. The chapter underscores sustainable practices, ensuring aluminium alloys remain essential for defense while advancing eco-friendly production.

Abstract: The flight operations inside the cockpit and fuselage are moving towards technological based systems, such as Tablets and Mobile computers. The new technology also offers a variety of possible options to pick the right tools and gadgets for regular operations performed by the Pilot, inside the Cockpit and in the fuselage. This paper surveys different gadgets, Electronic Flight Bag (EFB) used by the aviation pilots as well as Inflight Entertainment Devices (IFE) used in the aviation industry. Moreover, the present work also surveys the variety of gadgets used to achieve various goals with the existing issues and challenges.

Abstract: A curved fuselage C-frame manufactured using RTM has been studied under various curing conditions and the part deformations were measured.

Abstract: Within the Cluster of Excellence for Sustainable and Energy-Efficient Aviation SE 2 A, a blended wing body aircraft is investigated to improve efficiency and carbon emissions of future air transport. By embedding the aircraft engines on the top rear fuselage, parts of the aircraft’s wing boundary layer are ingested, which has the potential to further improve the engine’s propulsion efficiency. Through the ingestion of low momentum fluid, inflow distortion is induced and the fan rotor operates under increased flow incidence, when passing through the distorted flow regimes. To reduce the thereby arising efficiency and pressure ratio penalties in the aircraft engine, alternative design strategies for the fan stage are required. Within this investigation, an active shape morphing mechanism is introduced, which allows to temporarily adjust the fan blading when the fan rotor is exposed to distorted inflow conditions. By integrating piezoceramic actuators into the rotor blading, the blade staggering and turning can be adjusted with the goal to reduce flow incidence and deviation in the distorted flow regimes. For this investigation, the NASA rotor 67 is chosen as an initial test case and its performance under boundary layer ingestion (BLI) conditions is evaluated. For the shape morphing assessment, FEA morphing simulations are coupled with stationary CFD simulations of the actuated fan rotor geometries under distorted inflow. As the achievable deformations for the NASA rotor 67 are however too small to compensate for the strong distortion effects, a fan re-design is conducted. The re-design follows current ultra-high-bypass-ratio (UHBR) fan concepts with a particular focus on the shape-morphing capability of the rotor. Within this investigation the focus especially lies on three-dimensional design adaptions, such as a hub chord reduction as well as dihedral and sweep. By considering carbon fiber reinforced polymers (CFRP) as blade material, the impact of tailored blade architectures on the morphing behavior is additionally considered.

Abstract: The performance of a higher harmonic control system called the Multipoint Adaptive Vibration Suppression System (MA VSS) at reducing 3/rev wing vibratory loads and fuselage vibrations on a dynamically-scaled tiltrotor model is presented. Previous wind tunnel tests on a semispan aeroelastic tiltrotor model have demonstrated the effectiveness of MAVSS for reducing wing vibratory loads using both an active flaperon and swashplate. The primary goal, however, of such a vibration suppression system is to reduce tiltrotor fuselage vibrations in order to improve passenger comfort. The present study addresses the reduction of both wing and fuselage vibrations using MA VSS active flaperons on a 1/10-scale dynamic tiltrotor model designed to be representative of a tiltrotor configuration. Also, this study attempts to identify possible problems that may impede the application of MA VSS flaperon control forces for the purpose of tiltrotor vibration reduction in the presence of full-span symmetric and antisymmetric wing modes of vibration. Electromagnetic shakers applied simulated 3/rev vibratory hub loads and higher harmonic control forces, simulating active flaperons, to the tiltrotor model. MAVSS is shown to be effective at reducing individual and combined 3/rev vibratory wing loads in the semispan configuration. For the more complex full-span configuration, MAVSS flaperon control forces are shown to be effective at reducing wing vibratory loads or fuselage vertical accelerations, but not as effective at reducing wing loads and fuselage vibrations simultaneously. Vibration reduction trends are shown to be a function of the simulated rotor speed for both the semispan and full-span configurations. These results suggest that the application of MA VSS-controlled flaperons to a tiltrotor configuration may prove to be difficult due to elevated wing vibratory loads during reduction of fuselage vibrations and because wing vibratory loads and fuselage vibrations cannot be reduced simultaneously.

Abstract: Hail ice impact is a realistic and yet not completely understood threat to exposed composite structures such as aircraft fuselage and wing skins, leading edge surfaces, engine nacelles, and fan blades. To investigate this threat, experiments in which carbon/epoxy composite panels were impacted by ice spheres at high velocity (30 to 200 m/s) were conducted to measure: (i) the impact energy at which damage initiates, and (ii) elastic response of the composite panel resulting from impact. Subsequent numerical analyses were performed of the impacts and were validated through correlation with experimental data. Insights gained from the numerical analyses were used to compose an analytical formula predicting the onset of delamination. This formula, based on a global energy balance, provides a cost effective (i.e. lower number of tests needed) means by which the impact damage resistance of composite structures and of composite material types can be established.

Abstract: A thrust compounded helicopter—a main rotor with a trailing propeller was tested in the Glenn L. Martin Wind Tunnel to evaluate its performance under different flight conditions. The main rotor rig consists of a hingeless hub with four fully instrumented NACA 0012 blades and a modified Robin fuselage. The propeller rig consists of a rigid Sensenich L26H propeller with four blades. Tunnel tests were carried out for the isolated propeller and thrust compounded rotor configurations. The isolated propeller tests were conducted with and without the fuselage installed ahead of it to understand the effects of the fuselage on propeller performance. The thrust compound configuration was tested at three different main rotor shaft tilt angles (αs): −4°, 0°, and 4° (fwd tilt, vertical, rearward tilt), advance ratios (μ) from 0.3 to 0.6, and multiple lift (CL/σ) and propulsive (CX/σ) trim targets. Main rotor hub loads, oscillatory blade structural loads, and propeller hub loads were measured for all the tests. The test data were verified with a full vehicle aeromechanical analysis using the University of Maryland Advanced Rotorcraft Code. The thrust compound configuration with the main rotor shaft tilt of 4° (rearward tilt) provided the best performance. Thrust compounding with rearward shaft tilt (4°) resulted in a 50% increase in the maximum aircraft lift-to-drag ratio compared to a single-rotor helicopter. Half peak-to-peak hub vibratory loads and blade bending loads decreased with thrust compounding. It was observed that for the same lift target (CL/σ), thrust compounding achieved 20% higher flight speeds than a single rotor.

Abstract: With fuel economy being a significant factor in today's world of aviation, aircraft manufacturers have been looking into ways to decrease drag to save fuel. Drag reduction using riblets on wing and fuselage surfaces is . an area that has been studied extensively for the last two decades. In preparation for a larger project which will use several methods to simulate this phenomena, small scale research had to be done to find ways to generate similar conditions on scaled models in a low speed wind tunnel and examine boundary layer behavior. In this project, boundary layer measurement techniques, such as rake measurements and Preston tube measurements on a flat plate mounted on the wind tunnel side wall are presented with calibration techniques using "Design of Experiments" applied to electronic pressure scanners which were used to collect data. In addition, two methods to artificially thicken the boundary layer on a flat plate are presented in this thesis. The first technique is to use surface roughness elements and the second is to use a blowing technique in the wind tunnel. Measurements show that boundary layer parameters are in good agreement with theoretical flat plate results according to the rake and Preston tube measurements. The Preston tube technique proves to be an effective method to measure skin friction coefficient on a surface. Surface roughness elements indicate that the ability to thicken the boundary layer is feasible in a low speed wind tunnel. Artificial thickening by blowing can be applied in a wind tunnel, but the volume coefficient should be sufficiently high. Nevertheless, the techniques presented in this paper are shown to be feasible. Finally, a FLOPS analysis also demonstrated the potential performance advantages for actual aircraft assuming that skin friction reduction techniques were successful.

Abstract: This paper explores the practice of lightweighting in aerospace, which involves reducing a vehicle’s weight without compromising its functionality. Lightweighting is a rapidly advancing field with significant impact on commercial, private, and federal air/spacecraft. It is a top priority in aerospace research due to its broad benefits, including improved performance, reduced environmental impact, and increased operational efficiency. Additionally, lightweighting offers value to industries beyond aerospace, such as manufacturing and transportation. This paper delves into three main approaches to lightweighting: material innovations, structural optimization, and fuel advancements. Lightweight materials like alloys, composites, and polymers each have distinct advantages and drawbacks. Alloys are cost-effective and have a strong strength-to-weight ratio, while composites offer higher strength at lower weight but are more expensive and complex. Polymers can be tailored for specific functions, making them versatile in their applications. Structural optimization involves balancing durability and weight through advanced designs like ribbed fuselage shells and hybrid wing-body structures. Meta-heuristic algorithms help find optimal solutions to the structural problem of size, shape, and topology. Lastly, alternative fuel sources are explored for their energy density and potential to support lightweighting. Overcoming the challenges presented by new fuel sources and their characteristics proves an ongoing endeavor, however. Some radical and new ideas exist to reduce energy consumption uncorrelated with fuel source, also progressing lightweighting. Together, these methods contribute to significant advancements in aerospace technology.

Abstract: The fuselage serves as the primary component of commercial aircraft. The strength reliability of fuselage panels is therefore crucial for commercial aircraft. In the present study, a finite element (FE)-based modeling approach has been developed to predict the post-buckling behavior of curved fuselage panels under combined axial compression and in-plane shear loads at different shear-to-compression ratios. The intra-laminar damage was replicated using a progressive damage model driven by the Hashin’s failure criteria, while the skin−stiffener debonding was modeled using the cohesive zone model. Failure tests were performed using a bespoke Fuselage Panel Test System (FPTS), enabling comparison between experiments and simulations. The predicted buckling loads and ultimate failure loads are in good agreement with those obtained from experiments, which verify the predictive capability of the FE model. The failure load of the panels was found to be at least 30% higher than the initial buckling loads for all loading cases, indicating significant post-buckling load-carrying capacity. Under these four loading conditions studied, the load transfer mechanisms of curved panels were examined. All specimens experienced local skin buckling and subsequent global buckling, resulting in skin−stiffener debonding followed by fracture of the stiffeners, which was the dominant failure mechanism for the panel studied.

Abstract: Abstract When a seaplane lands in harsh sea conditions, it experiences substantial impact loads, which have a significant impact on the structural strength of the seaplane. This article explores the variation of the ratio of vertical acceleration to gravitational acceleration a z /g at the center of gravity of the seaplane model and the peak wave impact load on the bottom of the fuselage when the seaplane model lands in a wave environment with four different wave height-to-wavelength ratios h/l at three initial pitch angles through physical experiments. The experiment results show that the peak value of a z /g is larger when the initial pitch angle a = 6° compared to the initial pitch angles a = 4° and a = 8°, and the maximum value of a z /g decreases with the decrease of the wave height-wavelength ratio h/l . When the wave height-wavelength ratio h/l = 0.05, a z /g reaches a maximum value of 0.1265. At the initial pitch angle of a = 6°, the pressure peak at the bottom of the fuselage during the ditching shows an overall trend of first increasing and then decreasing from the nose to the tail, with a maximum value occurring near the 1/3 position of the fuselage near the nose.

Abstract: This thesis investigates the use of specially-designed “tension absorber” joints in composite vehicular structures for the absorption of energy in a crash situation through a process referred to here as “extended bearing failure”. The specific targeted application is future narrow-body composite aircraft fuselages which require an innovative energy absorption strategy due to the limited height available below the cargo floor for traditional crush beams. However, tension absorbers could be applied in any structure requiring energy-absorption capability in a crash or overload situation. Through a combined experimental-numerical approach, the work aims to provide fundamental information on the effects of geometric and material parameters such as stacking sequence, pin diameter, laminate thickness and loading rate, and an assessment of whether state-of-the-art numerical simulation is capable of providing genuinely predictive capability for such a complex problem. To make the results as useful as possible the chosen material is IM7/8552 carbon/epoxy, one of the most widely-characterised materials in the literature. Thus the results can be used by other researchers to test out modelling approaches without the need for further material testing. Besides the results in the published papers, videos provided as supplementary information contain complete three-dimensional (3D) maps of internal specimen damage, obtained from computed tomography (CT). The chosen performance parameters are ultimate bearing strength (UBS), mean crushing stress (MCS) and mass-specific energy absorption (SEA). Diameter-to-thickness (D/t) ratio is found to be an excellent predictor of UBS and SEA for both quasi-static and dynamic loading rates, with small D/t values giving best results, provided the thickness is sufficient to avoid global bending of the specimen. Concerning the effects of stacking sequence, it is found that the most important factor in maximising SEA is having small changes in orientation at ply interfaces. This is even more important than 0° content. Laminates with a high SEA tend to have a low UBS. Highest UBS was for quasi-isotropic laminates. Increased loading rate results in increased UBS but decreased SEA. The implemented model is a physically-based, three-dimensional damage model which uses in-situ ply strengths, stress-based fibre failure criteria, Puck’s criteria for matrix damage, a non-linear law for in-plane shear, a cohesive zone model for delamination, a crack-band model to mitigate mesh sensitivity, and frictional contact between the pin and the laminate, and between plies once they delaminate. The developed model is found to accurately predict the global response in terms of strength and energy absorption and can forecast the effects of changing geometry and material parameters. Critically, comparison with CT scans shows that it also captures the key mesoscale damage mechanisms.

Abstract: This paper presents a side-by-side comparison of most design concepts employed in the practice of sizing fuselage panels by the aerospace industry with a main function of structural weight.

Abstract: This paper presents modeling development for predicting progressive damage to composite aircraft fuselage-type panels subject to wide area blunt impact.

Abstract: Abstract The Blended Wing Body (BWB) aircraft is a non-conventional aerodynamic configuration that merges the fuselage and wings into a seamless structure, offering significant advantages in fuel efficiency, payload capacity, and aerodynamic efficiency. However, the absence of conventional tail units introduces aerodynamic and stability challenges that require careful design optimization. This study investigates the aerodynamic and static stability characteristics of a baseline BWB design from the European Distributed Multi-Disciplinary Design and Optimization (EU MOB) project. The analysis is performed using a mid-fidelity numerical tool based on the vortex lattice method, incorporating geometric and mass properties from previous studies. Aerodynamic curves and static stability derivatives are evaluated across a specific flight regime and compared with existing literature verifying the suitability of the used tool for preliminary aerodynamic and stability assessments. Subsequently, a parametric study is conducted to assess the effects of varying winglet design parameters, including height, sweep, cant, and toe angles, on aerodynamic efficiency and static stability characteristics. The results demonstrate that optimized winglet configurations can enhance lift-to-drag ratio and improve static stability characteristics without significantly increasing drag. These findings provide valuable insights into the role of winglets in improving BWB aircraft performance and contribute to the optimization of next-generation aerodynamic designs.