In recent years, aircraft noise has emerged as a pressing concern within aeronautics, due to its adverse health impacts and the increasing annoyance experienced by affected populations. Factors such as rapid urbanization, urban encroachment resulting in closer proximity of residential areas to airports, and the continuous growth of air traffic have exacerbated this issue. Current approaches to mitigate aircraft noise primarily focus on integrating additional components rather than optimizing the airfoil profile a priori.

This thesis aims to actively address this challenge by developing a coupled aerodynamic and aeroacoustic shape optimization framework tailored for a 2D airfoil. The primary objective is to minimize trailing edge noise, identified as one of the dominant noise generating mechanisms during approach and landing, while simultaneously maximizing aerodynamic performance. The solving strategy combines an aerodynamic solver with a state-of-the-art wall pressure spectrum model and Amiet’s trailing edge noise (TEN) model, whose inputs are boundary layer parameters extracted from the aerodynamic evaluation.

The efficacy of both lower fidelity (XFOIL) and higher fidelity (Reynolds-Averaged Navier Stokes (RANS)) aerodynamic solvers is evaluated. The airfoil is optimized at different points of the flight envelope: for maximum lift-to-drag ratio during cruise and for minimum trailing edge noise during landing.

Results demonstrate that the genetic algorithm (NSGA-II) optimization framework yields promising airfoil shapes and reliable outcomes at a computationally feasible cost. Many optimizations with varying generations and population sizes are ran: remarkably, the results consistently showcase well-defined Pareto fronts, with superior definition observed particularly at a population size of 200.

The lower fidelity solver proves particularly effective in the landing scenario, while it shows limitations for higher Mach numbers. In this view, the RANS code is necessary for capturing flow phenomena at cruise speeds, and notably, it also provides convincing results in terms of aeroacoustic prediction during landing phase.

In general, this research highlights the potential for significant advancements in designing optimal airfoils for aerodynamic performance and TEN.

The implementation of morphing wings technology has the potential to significantly improve aircraft performance. However, due to the complex interplay between aerodynamics and structural integrity, it requires an aerostructural analysis. 3DExperience by Dassault Systèmes is a Computer-Aided Design (CAD) - based software with integrated structural and aerodynamic analysis tools. This software was employed to establish a methodology for conducting one-way aerostructural analysis of morphing trailing edge wings on a transonic transport aircraft. The methodology highlights the benefits of CAD-based modelling and the coupling between the aerodynamic and structural models. Nevertheless, the aerodynamic analysis results were affected by a node limit and the meshing options. Finally, a high-lift and an aerodynamically efficient test case were utilised to compare the impact of the trailing edge morphing to that of an unmorphed wing.

This thesis presents the development of an entropy-corrected full potential flow solver (EC-Flow) within a finite element framework to enhance the representation of shock waves in transonic flow regimes. By addressing limitations in traditional isentropic full potential flow solvers, EC-Flow improves aerodynamic predictions whilst maintaining low computational cost, making it suitable for early design phases in aeroelastic analysis.

The methodology integrates an entropy correction based on the Rankine-Hugoniot relations, with shock detection achieved through gradient-based and normal Mach number methods. An iterative downstream approach ensures consistent application of the correction across the shock-affected domain.

Results for the NACA0012 and RAE2822 airfoils demonstrate improved shock representation, and faster convergence compared to the uncorrected solver. Validation against Euler solutions confirms the effectiveness of the entropy correction in enhancing accuracy without significant computational overhead.

In the transition from fossil fuels to renewable energy sources, the advancement and deployment of wind turbine technology plays a crucial role. A notable trend in wind turbine design is the ever-growing rotor size, which entails that wind turbine blades have become very slender and flexible structures.

This increasing flexibility offers an opportunity for tailoring the aeroelastic behaviour of wind turbine blades. One such aeroelastic tailoring technique is blade sweep, defined as a displacement of the blade axis in the rotor plane. Blade sweep couples bending and torsion deformations and can thus be used to passively alleviate loads on the blade.

For the design, optimisation, and certification of wind turbine blades, blade element momentum theory (BEM) remains the aerodynamic simulation method most relied upon. However, BEM-based numerical tools inherently assume a straight blade geometry and, hence, cannot accurately model the additional flow complexities introduced by blade sweep.

This dissertation starts by presenting a newly developed BEM correction model for swept blades. The focus is on accurately modelling the azimuthal displacement of trailed vorticity and the curved bound vortex self-induction while maintaining BEM's streamtube-independent approach and rapid calculation speed. The developed model shows good agreement with mid-fidelity modelling (lifting line simulations), which intrinsically can model the two aforementioned effects of blade sweep.

To validate the BEM correction model for swept blades beyond the comparison with lifting line simulations, two wind tunnel campaigns are conducted, one with straight blades, being thrust-scaled versions of the IEA 15 MW reference wind turbine blades, and one with swept blades. While the former is intended to provide a baseline for the accuracy of numerical modelling, the latter then provides means to assess the impact of blade sweep and how this is captured in low-fidelity numerical simulations. The validation is conducted based on blade-aerodynamic quantities derived from flow fields measured using particle image velocimetry (PIV). It is demonstrated that the application of the BEM correction model improves the match with the experimental data compared to simulations without the correction model being applied.

Furthermore, this dissertation covers three diverse research efforts conducted within the framework of the TIADE project, a field experiment on a full-scale wind turbine. Firstly, a robust approach to optimise the spacing of pressure sensors for aerodynamic measurements on wind turbine airfoils is presented. The approach considers the expected turbine operating conditions and improves the lift prediction accuracy compared to a simpler, cosine sensor spacing over a wide range of angles of attack. Given that two fundamentally different optimisation routines arrive at close-to-identical solutions, it can be concluded that an optimal solution exists for placing pressure sensors around an airfoil to conduct aerodynamic measurements.

Secondly, pressure measurements obtained on the TIADE research wind turbine over multiple months are employed to validate aeroelastic simulations. The validation is performed based on both ten-minute average data and time-resolved data and using both the integrated sectional forces and the underlying pressure distributions. Generally, a reasonably good agreement between simulated and measured data is found. This indicates that BEM-based aeroelastic algorithms are still valid tools to simulate modern, multi-megawatt wind turbines and their slender and flexible blades.

Finally, a design study of a blade with swept tip for the TIADE field experiment and thus under realistic geometric and load restrictions is conducted. Simulations suggest that flapwise fatigue and extreme blade root loads can be reduced. The same holds for the fore-aft and yawing moments at the turbine tower base. Simultaneously, the turbine performance in terms of power output remains unaffected. These results highlight the potential benefits of blade sweep as an alternative tip geometry for modular blades or as a conscious design choice for future generations of blades.

In conclusion, this dissertation contributes to a more accurate understanding and numerical modelling of swept blade aerodynamics. By moving from fundamental analyses all the way to more applied investigations of swept blade tips for a field experiment, the presented research helps pave the way towards swept blades being a valid option in future wind turbine designs.@en

The flared folding wingtip (FFWT) is a concept presented by Airbus to increase the aerodynamic efficiency and provide a means of load alleviation. A gate-to-gate demonstration of this concept has already been presented with the AlbatrossONE, a scaled aircraft featuring the FFWT concept. In this demonstration, the aircraft fold the wingtips during taxiing to fulfil gate requirements and extends it before take-off for improved efficiency. In addition, they can be released during flight to provide load alleviation and reduce the roll rate reduction caused by the increase in the wingspan. The objective of this thesis is to further study the FFWT concept.

An aeroelastic wind tunnel experiment to identify the influence of the wing stiffness and hinge release threshold on the gust load alleviation performance of a folding wingtip design is presented in this study. Five models with different stiffness and tailoring properties are tested and the wing root bending moment at different conditions is compared to the response with locked hinge conditions to assess the impact on the gust load alleviation capabilities of the folding wingtip.

The results show that the structural properties do not have an important impact on the peak load alleviation but the hinge release threshold and timing do. Releasing with the correct timing can reduce significantly the peak loads. However, the dynamics of the system are affected by this release: the flutter speed is decreased and, although the performance can improve, load oscillations increase, which can be considered detrimental for reasons such as fatigue.

With the increased focus on sustainability, aircraft are designed to reduce their emissions. One way to accomplish this is by increasing the wing aspect ratio, thereby increasing the aerodynamic efficiency, however this is not without consequences. Increased aspect ratio wings have a higher structural mass, are more susceptible to gust and maneuver loads and generally flutter at lower velocities. Due to the advent of both passive and active control techniques these issues can fortunately be solved by using gust load allevation (GLA), maneuver load alleviation (MLA) and flutter suppression.
Current aeroelastic testing facilities at Delft University of Technology include a gust generator and aeroelastic apparatus, used to suspend a passive wing section in the wind tunnel. The need for ad¬ ditional research on aeroelastic control in order to improve the sustainability and safety of aviation necessitates the development of a new wing section with aileron and spoiler control surfaces that is compatible with current facilities. The development, manufacturing and initial characterization and test¬ ing of this wing section is the subject of the present work.
As the new wing section includes a spoiler, a literature review is performed on this subject. Spoilers function by deflecting into the flow, causing separation aft of the spoiler and creating a large turbulent wake, resulting in a drastic decrease of lift. A linear potential flow model for spoiler aerodynamics developed by Brown and Parkinson was implemented in MATLAB with the intent of implementing this in future aeroelastic models. Verification of this model showed good agreements with original data presented in the paper describing the model.
The passive wing section was chosen as a basis for the new design. The position and size of the control surfaces are determined based on a review of experimental and operational applications. The new wing section was designed, resulting in a self¬contained model, including a single¬board computer, sensors and power supply. Actuation mechanisms were developed for the control surfaces, with a parametric device for control surface free play included in the aileron actuation mechanism. The new wing section was manufactured successfully and control software was implemented using Simulink.
A series of tests were performed to characterize the dynamic behavior of the wing section. Due to a combination of higher inertia and kinematics of the actuation mechanism, the usable bandwidth of the aileron is shown to be lower than that of the spoiler. Aerodynamic results show that the combined use of aileron and spoiler result in a reduction or reversal of the aerodynamic response of the wing. Gust load alleviation results with proportional control show an increase in damping by 1300% and a reduction in peak amplitude of 50% when using the spoiler. Results for the aileron are notice¬ ably less, with a decrease in amplitude of 15% and an increase of damping of 145%. The differences are attributed to both the differences in kinematics of the mechanisms as well as the greater absolute change in lift coefficient obtainable by the spoiler.

With the increasing desire of the aerospace industry to reduce emissions and fuel consumption, morphing wings have gained much interest due to the ability to adapt the wing shape in-flight for improved energy efficiency and aerodynamic performance. Active wing morphing is a technology that can improve aerodynamic performance continuously through different flight phases. However, a multidisciplinary approach is needed, which integrates the design, modelling, sensing and control methodologies in a multi-objective framework, and allows the smart autonomous morphing wing system to adapt its shape autonomously.

The SmartX project was initiated for this purpose at the Delft University of Technology, Faculty of Aerospace Engineering, Department of Aerospace Structures and Materials, aiming to investigate the energy-efficient wing concepts through smart wings.

This dissertation presents the Development, Realisation & Validation of a smart morphing wing, the SmartX-Alpha, capable of meeting various real-time objectives with distributed seamless morphing modules. This is done through a holistic approach considering all building blocks of a morphing system presented in four Parts of the dissertation.

Part I tackles the sensing approach required to reconstruct the shape of the wing in real-time with a vision-based sensing approach. Part II presents the design, development, realisation and experimental testing of a distributed modular morphing concept, SmartX-Alpha. Part III presents the multi-objective control framework developed to meet the gust and manoeuvre load alleviation objective and the real-time shape optimisation strategy to improve online aerodynamic performance. Furthermore, a vision-based control strategy is proposed to mitigate nonlinearities in the actuation system arising from mechanical imperfections. A series of wind tunnel experiments are conducted in the OJF to validate the methodologies on the SmartX-Alpha, ensuring the objectives are satisfied autonomously, in-real time. The final Part, Part IV presents the development of a second wing demonstrator, the SmartX-Neo, with distributed discretised control surfaces incorporating the previous learnings.@en

High Altitude Long Endurance (HALE) configurations have gained much popularity in the recent past. Although, HALE aircraft enable efficient and environmental-friendly flying, they necessitate a large Aspect-Ratio wing. These slender structures are prone to large, non-linear deformations which further increase the unsteady nature of aerodynamic flow that can detrimentally impact the structure. In the light of this trend towards an efficient aviation industry of tomorrow, Airbus Innovations UK has ventured into the development of a folding wing tip concept whose benefits are two-fold. First, is the ability to increase span during flight, thereby reducing the induced drag. The folding capability also ensures entry gate requirements at airports. Secondly, the folding mechanism presents load alleviation capabilities with appropriate hinge setup. Alleviation of loads on lifting surfaces directly translates to reduction of the Wing Root Bending Moment (WRBM), which means a reduction of weight. Thus, a further improvement in efficiency of flight. In this work, a nonlinear aeroelastic solver is developed. The proposed partitioned framework couples the industry standard commercial software package MSC Nastran, a structural solver to the Unsteady Vortex Lattice Method (UVLM), an aerodynamic solver. The UVLM module developed, exhibits both static and dynamic capabilities. Three wake models are incorporated. Horse-shoe wake is employed for static simulations and convective wake procedures are available for transient dynamic simulations. The independent modular solvers are coupled using Radial Basis Functions (RBFs) and Nearest Neighbour (NN) approaches. This framework enables the solver to capture the combined effects of structural nonlinearities and unsteady flows. The proposed solver is then verified and validated for two different test cases: NASA CRM Wing and Pazy Wing. Experimental validation is carried out using existing results from related work. The solver is then used to evaluate the “AlbatrossOne”, a hinged wingtips concept by Airbus Innovations UK. Both static and dynamic responses are analysed for load alleviation capabilities.

This work is a part of the Automated Rotor Blade Inspection (ARBI) project. ARBI aims to improve the rotorcraft RTB process with novel rotor blade measurement equipment such as 3D scanning, thermography and shearography. The research aims to quantify the effects of changes in blade properties on rotor vibration levels for CH-47 Chinook. This is done in order to create a rough ranking of which parameters have the highest impact on vibrations. This was performed by changing the blade properties on a single blade in the rotorcraft simulation tool FLIGHTLAB, trimming to a desired flight condition and collecting acceleration measurements. In FLIGHTLAB a novel inflow solver, the Viscous Vortex Particle Method (VPM) was applied, since the developers specified it was well suited for rotorcraft with rotor-on-rotor aerodynamic interaction. The results of this research in future hopefully will be able to help identify blades with poor RTB potential and help with suggesting blade sets with higher RTB success chance based on measurements from ARBI.

With the focus on the reduction of fossil fuel emissions, aircraft are continuously growing towards higher fuel efficiency. The traditional limits of aircraft performance can be surpassed through the use of composite materials which offer a reduction in aircraft weight. Due to the multidisciplinary nature of aircraft design, integration of different disciplinary analyses is required to arrive at a feasible design. The inclusion of composite design in the preliminary design process, however has become a challenge. This is due to the high computational cost associated with the composite aeroelastic tailoring tools used in the design process. A possible solution is available in the form of surrogate models which can reduce the computational costs. The current work focuses on the development of a methodology that allows the inclusion of surrogate-based model in the optimization for a computationally expensive aeroelastic tool (PROTEUS) developed at TU Delft. The resulting methodology can be expanded to a generic, computationally expensive tool in a multidisciplinary optimization setting. Wing design optimization is carried out based on surrogate modeling methodology and metal based design method. Comparison is made of the final optimized designs based on structural and performance parameters.

Electrification of aviation is driving new aircraft configurations consisting in slender, lighter wings with several propellers spread across the wing span, including the wing tips. Thinner wings with high aspect ratios tend to be more flexible, and thus, more susceptible to wing flutter, a dynamic aeroelastic instability characterised by diverging oscillatory motions of the wing. As propellers are flexibly attached to the wing, another aeroelastic instability, called precession or whirl flutter, can occur. The nature of this phenomena lies in the additional forces and moments created by rotating blades as the propeller hub describes a “whirling” motion about the static thrust axis due to gyroscopic effects. These dynamic instabilities can happen independently or coupled together causing severe structural damage on the aircraft and when occurred during flight, lead to fatal accidents. Hence, it becomes essential to analyse any new propeller-driven aircraft configuration regarding wing flutter and whirl flutter. A mathematical model has been derived to analyse the dynamic behaviour of a flexible cantilever wing with flexibly attached tractor propeller/s: the wing structural model is a linear dynamic Euler-Bernoulli beam model, the wing aerodynamic model is represented by a combination of strip theory and two-dimensional Theodorsen's unsteady aerodynamic theory, and the propeller aerodynamic model is defined by Houbolt-Reed propeller aerodynamic derivatives. The aeroelastic behaviour of a wing with flexibly mounted propeller/s may be very different from that of a wing with rigidly mounted propeller/s. Reasons lie in the interaction between the additional whirl modes and the wing modes due to gyroscopic coupling. Whereas propeller and wing aerodynamics may drive the (coupled) backward whirl mode unstable, wing aerodynamics alone may drive the (coupled) forward whirl mode unstable.

Classical aerodynamic wind tunnel tests are performed with stiff or rigid wind tunnel models. Over the years wing structures have become more flexible, and with the rise of non-conventional designs, such as strut-braced wings featuring high aspect ratios, even more flexible designs have risen. To capture the aerodynamic performance of these concepts under varying load conditions, resulting in different flight shapes, a static aeroelastically scaled wind tunnel model is required. The thesis research objective is to structurally design a static-aeroelastically scaled wind tunnel test model using optimization routines to find the optimal scaled aeroelastic characteristics representing the reference wing in wind tunnel test conditions while incorporating the safety constraints. Use cases included in the research are both an unconventional high aspect ratio strut-braced wing as well as a highly flexible cantilever wing.

To obtain aeroelastic and aerodynamic similarity, the aerodynamic similarity parameters should be the same and the elastic properties of the structure should adhere to static-aeroelastic scaling laws, which were found to solely depend on the length and pressure ratios. For finding a structure that has those elastic properties, design optimization is needed. After length scaling for the wind tunnel, little room for structural elements is left, especially in the slender strut-braced wing concept. For this reason, a composite shell with solid foam is used, as composites allow control over the directional stiffness. However, the discrete nature of composites, consisting of various layers with each a variable ply direction, complicates the optimization. To overcome this complication, lamination parameters are used to make the design space continuous, allowing the use of gradient-based solvers. An optimization framework was developed to optimize the lamination parameter space including a finite element solver in combination with scaled aeroelastic loads. The design variables considered in the optimization are the lamination parameters and the thicknesses of the laminate with the root mean square of the difference in deformation as objective. After optimizing in the lamination parameter space, the stacking sequence is retrieved using an open-source tool, OptiBless.

The results showed that the design space of the unconventional strut-braced wing configuration is highly non-convex which results in finding local minimum solutions. For a cantilever wing, the optimization is much more robust, as the optimization yields comparable results for multiple starting points which can be considered as the global optimum. The optimization framework developed was able to design a static aeroelastically scaled wind tunnel model, consistently resulting in a representative match in terms of deformation with the full-scale design for the cantilever wing. For the strut-braced wing, a representative match can also be found but these results were inconsistent. The optimization framework proved to be successful to optimize composite designs for cantilever wings. For non-conventional designs, this method is less powerful as using a great number of starting values is required, diminishing the computational benefit of using a gradient-based solver, and other strategies could be considered that may yield better results.

The next step in spaceflight, among others, is the development of a reusable launch vehicle (RLV) suitable for manned missions. All existing designs of re-entry vehicles are either ballistic or low L/D vehicles. In the future, the aim is to develop a high L/D vehicle, because such vehicles experience lower g-loads, making the flight comfortable for manned flights. The development of these vehicles can be useful for both, space (for example, space tourism) and military applications (for example, long-range missiles). Thermal protection system (TPS) is deemed critical to the RLV development, as high L/D vehicles are expected to experience higher thermal loads. Existing TPS solutions are not suitable for this purpose. Therefore, arises the need to find reusable TPS solutions that can sustain the desired thermal loads. Flight testing is a crucial step for developing any hypersonic system, this applies to the TPS design as well. This study aims to investigate the influence of TPS on designing the mission and vehicle for a test flight.
From the various developed and proposed TPS designs in literature, a TPS solution is selected, to meet the requirements identified for future RLVs. The solution is an active cooling concept (cooled metallic TPS), named Enhanced radiation cooling, that uses water as a coolant. Based on a preliminary design investigation, suitable materials are selected and the thickness of the layer is determined. The performance of the proposed TPS design, i.e., the thermal load-bearing ability, is assessed using a transient thermal analysis tool developed for this purpose. The outcome of this analysis proves that the proposed design does not meet the thermal load requirements identified in this study. Nonetheless, the concept shows substantial improvement in performance, it can sustain almost double the heat flux as compared to an uncooled system.
To meet the heat flux requirements, a few modifications in the design are proposed and analysed. These modifications increase the complexity of the system and have other adverse consequences that must be addressed. However, thermal analysis proves that the heat flux requirements, identified at the start of this study, are satisfied. A sensitivity analysis of this modified cooling system shows that the performance of the design is not influenced by material uncertainties. Additionally, the design is found to be robust under varying mission and design parameters, within reasonable bounds. These results can serve as a preliminary input for designing a test vehicle and mission.

This thesis embarked upon establishing and validating a design process to build and control a distributed TRIC concept seamless smart morphing wing to achieve simultaneous load alleviation, flutter suppression and drag minimization capabilities. A novel aeroelastic simulation tool was built to carry out composite wing skin optimization, with prototype testing carried out to design a flexible connection between modules. The work output was a fully built morphing wing which was successfully tested in the TU Delft Open Jet Facility. Control of the wing was partially made feasible through the development of a surrogate model of the system which was also validated with DIC testing.

Multi-step optimization of large composite structures involves formulating the stiffness distribution in terms of lamination parameters that provide a continuous design space, followed by a second optimization in the discrete design space to obtain equivalent stacking sequences. The retrieval of stacking sequences is often challenging as a result of additional manufacturing constraints such as blending in the discrete design space which are not factored into the initial continuous optimization. By subjecting the continuous optimization to blending constraints, a better match between the discrete and continuous design spaces is achieved which in turn improves the retrieval of stacking sequences with equivalent stiffness and mechanical behaviour.
An open-source Python application capable of multi-component optimization was developed in this thesis in order to revisit the application of lamination parameter matching objective functions in the discrete step. By subjecting these objectives to mechanical constraint penalties, the vicinity of reference continuous optimum could be explored more effectively to steer the discrete solutions towards feasible design regions. A comparison was made between discrete optimization according to blended and unblended continuous designs points, where it was found that the former translated to improvements in the retrieved stacking sequences with a considerable decrease in weight versus target continuous references without blending constraints of an aircraft wing.

Recent decades have seen the range of applications for aircraft expand to niches like weather monitoring, reconnaissance and satellite launch (Air Launch to Orbit, ALO). The requirements for such roles necessitate unconventional features like twin fuselages (in case of ALO mother ships) and very high aspect ratio wings (for long endurance missions). For conventional aircraft, aside from improving safety, the driving factors for new designs have always been increased efficiency and performance. The prevalent approach has been to reduce weight and induced drag by using higher aspect ratio wings. Consequently, newer designs are lighter, more flexible and closer the failure limit. Therefore, the interaction between flight loads and the airframe, i.e.\ aeroelasticity has become important for a safe design. High aspect ratio wings undergo large magnitude, low-strain deformations. An accurate understanding of their aeroelastic behaviour requires nonlinear analysis methods. This thesis implements a reliable nonlinear aeroelastic analysis method in NASTRAN. The aeroelastic analysis modules in commercial finite-element analysis software are aerodynamically and structurally linear. At the same time, such software often has advanced nonlinear structural analysis capabilities. This thesis modifies and combines two approaches mentioned in literature. NASTRAN's aeroelastic module is used to obtain rigid aerodynamic loads and its nonlinear structural module to obtain structural deformations. For a given angle of attack, altitude and airspeed a wing is analysed iteratively until the rigid air loads and structural deformations for successive iterations converge. The nonlinear structural module is used to obtain pre-stressed structural modes for this converged condition. These modes are then used in flutter analysis. An idealised HARW with arbitrary properties is analysed using this method and using TU Delft's in-house aeroelastic optimisation software: Proteus. Static aeroelastic deformations and flutter analysis results are compared for validation. Following this, the results of the linear and nonlinear aeroelastic analysis are compared. It is shown that linear analysis over-predicts the deformation in flexible wings and that the inclusion of pre-stress changes the damping of critical flutter modes; thereby changing the flutter point for the wing. For the test case analysed in this thesis, the inclusion of geometric nonlinearity and follower force effects resulted in large changes in the flutter speed with change in the angle of attack. In this way, the thesis successfully implements a nonlinear aeroelastic analysis method which can be used to improve aircraft designs during the preliminary stage.

Fundamental phase in the preliminary design of an aircraft is the load case choice. Historically, the manoeuvres have been considered as the principal critical loads, while in conceptual design the gusts have been not investigated since, being an intrinsically dynamic phenomenon, they are expensive to analyze. Nevertheless, they are a critical certification requirement which is becoming more and more demanding with the introduction of composite materials and high aspect-ratio wings. There is, however, an absence of methodical ways to take decisions regarding including the gust loads into the optimization in conceptual design. This MSc thesis aims to define design guidelines for gust loads inclusion in aeroelastic optimization for the civil transport aircraft category. Several optimization cases are computed in the TU Delft in-house aeroelastic framework PROTEUS, varying main design parameters, materials and gust models to understand the sensitivity of the design to the gust response analysis.

Morphing winglets are devices that can improve aircraft aerodynamic efficiency and also be used as load alleviation devices. However, this implementation comes with a weight penalty due to the morphing mechanisms. In the present study two morphing winglet cases are tested, where due to mechanisms compatibilities, only a variable Cant-Span winglet and a variable Sweep-Span winglet are proposed. For these two cases, the cant angle mechanism presents a high actuator requirement of about 750 [kg], when restricted to a winglet AR of 3.16 and a minimum cant angle of 33 [deg]. For the morphing sweep case, the weight penalty is of less than 50 [kg]. As a result, during a mission analysis, the extra weight penalty for the first winglet case increments the MTOM of the aircraft in 2500 [kg], while for the second case, the MTOM is reduced by about 1000 [kg]. The consequence of this is a higher fuel consumption for the first case, but a lower fuel consumption for the second one, where a fuel reduction of a 2.2% is achieved. Another benefit is a lighter wing structure due to the load alleviation capabilities, and a higher range of almost a 9% for maximum fuel is obtained.

An approach to optimizing variable stiffness laminates for strength is modified to allow for the optimization of curved structures. A density function to take in account the effect of gaps and overlaps during optimization created by fibre steering is found with the functional approximation method. The density plot of a plate for 2 standard cases is compared, after encountering complications an alternative solution is proposed with the finite element method.