A rotary-wing aircraft, or helicopter, is crucial for maritime operations due to its Vertical Take-Off and Landing (VTOL) capabilities, making it essential for small-deck Naval ships. Helicopters serve multiple roles, enhancing the Royal Netherlands Navy’s (RNLN) missions. However, the rise of Unmanned Aerial Vehicles (UAVs) in the past decade has introduced a cheaper, expendable, and more efficient alternative for certain operations, as UAVs require no pilot and can operate continuously. UAVs are ideal for tasks related to improving situational awareness, thereby increasing mission effectiveness, reducing pilot risk, and lowering fuel consumption. The RNLN aims to utilize UAVs alongside crewed helicopters on new ships to maximize operational efficiency. Both helicopters and UAVs follow three mission phases: launch, mission, and recovery, with the recovery phase being the most challenging. During the recovery phase, a crewed helicopter approaches the sailing vessel in a standardized way for which operational limits are established. These limits are essential for rough weather situations and are determined through many flight tests, which is a time-consuming, expensive and hazardous task. In addition, these operational limits only apply to a standardized approach manoeuvre. No approaches for UAVs are prescribed yet, making it difficult to determine such limits, but opens the possibility to standardize newly optimal UAV-ship approaches for which operational limits can be established. Moreover, the ship needs to manoeuvre to obtain optimal wind conditions for an approaching helicopter, which is not desired for UAVs. This means that UAVs should be able to approach and land on the ship in any wind condition. This study investigated the effect of wind on optimized UAV-ship approaches, which can be used to aid and standardize maritime UAV-ship approaches. For this objective, a trajectory optimization framework was created that combined helicopter dynamics and performance, a ship, three wind models of increasing levels of accuracy and a solver to find optimal wind-affected approaches. Trajectories were optimized to be smooth and fast because this is most important for operational purposes. Experiments involved several wind directions, wind speeds and helicopter starting positions, affected by three wind models. This work shows that the created framework functions properly and can be used in a variety of situations. The optimized trajectories were evaluated by the Longest Common SubSequence (LCSS) similarity measure to investigate the effect of both wind direction, speed and model fidelity. With an output between 0 and 1, the LCSS algorithm provided intuitive results and clear trends. A conclusion is that the helicopter wants to exploit the wind as much as possible. It diverts its path to acquire a higher ground speed at the start and a stronger headwind to decelerate at the end of the approach. When introducing a wind gradient, it was observed that the helicopter adjusted its vertical manoeuvre with the same objective. In some wind cases, constraint limits were reached or violated, marking preliminary operational limits. Results were also interpreted with a different perspective to find the best starting positions for the helicopter to approach the vessel in a certain wind condition. These results showed a variety of optimal starting locations, based on the total objective function value, total energy consumption and by distance normalized quantities. Wind model fidelity did not influence this outcome significantly, but it showed that it is important to estimate performance parameters with higher-fidelity models for increased wind speeds. This could also be seen from the downdraft behind the hangar; while the optimized approaches did not deviate geometrically, performance indicators such as the power required were affected significantly.

A propeller has a high propulsive efficiency, yet inherently leaves the air it propels swirling. The efficiency of a propeller can be enhanced by placing it in the tip vortex of a wing. Such a tip-mounted propeller configuration lets the vortex created by the wing and the vortex created by the propeller counteract each other, enhancing the propulsive efficiency of the system. Even though the benefits of such a configuration are clear, the concept is not commercially viable. Among other benefits, electric engines are scalable, light and can be placed in advantageous positions. This allows the engine to be placed at the tip of a wing. The batteries for electric propulsion are heavy, so synergistic effects are needed to improve the efficiency to such an extent that electric propulsion will be commercially viable. Another benefit can be defined by tip-mounted propulsion. As the propeller is mounted at the tip of the wing, the moment arm is large. This could result in a large directional control moment. On the other hand, if one engine is inoperative, the resulting moment is so large that it cannot be compensated by a rudder deflection. Therefore, if a one-engine inoperative situation arises, to still produce thrust both propellers have to be interlinked resulting in a heavy system, or both engines should be switched off. If tip-mounted propellers are used to enhance directional stability and control, the vertical tailplane size can be reduced, resulting in a reduction of mass, drag and hence overall power consumption. In this thesis the contribution of tip-mounted propellers to static stability, dynamic stability, and control is researched, for both positive and negative input power. Negative input power denotes that the propellers are recuperating energy fromthe airflow, and is a state that is applicable when one engine is inoperable. The forces on a tip-mounted pusher propeller are obtained by linking a lifting line wing model and a combined blade-element momentum vortex propeller model. The lifting line model is used to calculate wing-induced velocities on the pusher propeller disk. These induced velocities are used as input for the propeller model, and a resulting thrust and power is obtained. These are corrected for an angle of attack on the propeller disk by empirical relations. The thrust, normal force and power are saved to a 7D dataset. A non-linear flight mechanics model of the Piper Seneca III uses the dataset as lookup table and implements the tip-mounted propeller as forces at the tip locations. Thrust variations in the 7D dataset are parametrically visualised, showing the expected trends. Resulting forces are compared with a rudder deflection, to estimate a potential rudder size reduction. When the aircraft is flying slow, tip-mounted propellers can match the maximum moment produced by the conventional rudder. When flying fast, around half the reference yawing moment can be produced by tip-mounted propellers. The static stability contribution of tip-mounted propellers is visualised parametrically for propeller diameter, advance ratio, blade pitch and a toe-in angle. Negative toe-in angles, hence toe-out angles, prove to greatly enhance the directional static stability. High thrust values greatly enhance longitudinal static stability as the reference aircraft’s wing tips are positioned above the centre of gravity. In the same parametric fashion as the static stability evaluation, the contribution of tip-mounted propellers to dynamic stability is evaluated. The damping factor and natural frequency are obtained for first order motions Dutch roll and the Phugoid by fitting an exponential curve to the time-response of the flight mechanicsmodel. This method succeeds in capturing the motion characteristics, and indicates clear trends: the frequency of the Dutch roll increases with toe-out angles and thrust, and the frequency increases with thrust. The Phugoid’s natural frequency increases with thrust, and damping factor decreases with thrust. An attempt is done to use the curve-fitting method for the second order motion Short period as well, yet proves too inaccurate. A typical tip-mounted propeller design is evaluated to summarise the effects of using tip-mounted propellers for directional stability and control. For this non-optimised design, the static stability is enhanced by 37%. The dynamic stability has increased significantly as well. Maximum blade pitch deflection and maximum rudder deflection result in similar time-history sideslip responses when the aircraft is flying slow, when flying fast the tip-mounted propeller design’s sideslip response is around half the reference rudder deflection response. This conclusion also applies to a one engine inoperative simulation. In this thesis, a significant first step is made toward enhancing directional stability and control with tip-mounted propellers by defining the contribution of tip-mounted propellers in a parametric fashion. The results are directly applicable to an aircraft that has been designed with tip-mounted propellers, yet does not rely on tip-mounted propellers for directional stability and control. A typical design is evaluated in cruise, recuperative and one-engine inoperative cases, proving that from a stability and control point of view the vertical tailplane can be reduced in size. Before tip-mounted propellers can be used for directional stability and control behaviour analysis in the stalled regime, more research has to be conducted. This is left as recommendation.

With increasing fuel prices, airlines are searching for means to reduce fuel consumption. Due to the limited variables that operators can influence of an aircraft, they are left with only operational changes, such as center of gravity positioning. By shifting the center of gravity aft, achieved by loading heavy cargo and baggage in the rear, trim drag will decrease. Despite the relation between the center of gravity and the trim drag, and therefore fuel consumption being known, no literature exists on its quantification in practice. As such, at the request of KLM Loadcontrol, a numerical model was developed to investigate the effect of the center of gravity on fuel burn. Due to the complexity, the results were used to derive a function per aircraft, route and season. Defining the zero fuel weight and the center of gravity location for that condition will enable users to determine the effect.

Aircraft manufacturers are increasingly exploring emission-free flight or emission reduction for larger passenger aircraft. The low energy density of state-of-the-art battery technology limits the application to small, electric, fixed wing aircraft up to a flight time of approximately one hour. To overcome these limits, a combination of fuel cells and batteries to exploit the benefits of battery power density and hydrogen energy density was studied. Current Lithium-Ion battery cells reach approximately 1.6 kW/kg of maximum power density, much higher than fuel cell systems. On the other hand, the energy storage capacity of suitable hydrogen storage methods is much larger than battery cells, the latter have an energy density of 240 Wh/kg.

Because most demonstrated applications are for fixed wing aircraft, the unmanned GeoCopter GC-201 helicopter was used for performance requirements, weight and volume analysis. The study focuses on the preliminary sizing of the powertrain and the optimization of fuel cell and mission profile variables for this vehicle. Helicopter performance modelling, fuel cell static behavior as well as a battery discharge simulation are combined with lower fidelity models for other components.

The study results in a comparison of battery-only and fuel cell-battery configurations through payload-range diagrams, allowing for a quick evaluation of application areas. These mainly show that batteries excel at high payload, low range applications whereas a fuel cell-battery combination shows clear advantages at low payload, longer range applications. Liquid hydrogen will be shown to be comparable to the current micro gas turbine powered rotorcraft, with 400 and 500 km range capabilities respectively. Range capabilities for 300 bar and 700 bar compressed gas tank storage options show 140 and 180 km, with battery-only reaching a maximum range of 80 km.

New generation combat aircraft are expected to operate over extended flight envelopes, including flight at high flow angles and rapid maneuvers. Conditions beyond traditional limits are giving rise to nonlinear phenomena, such as flow separation, large scale energetic vortices, fluctuations etc. These phenomena have significant impact on aircraft performance and if not resolved accurately design uncertainties are increased risking lack of performance or even costly redesigns. Thus, accurate modelling of unsteady nonlinear aerodynamics is essential for modern and future combat aircraft.

Unfortunately, conventional modelling tools either lack the required fidelity or they are too expensive. Traditional, highly-efficient approaches are not suitable for modelling nonlinear flow phenomena. Concurrently, high fidelity Computational Fluid Dynamics (CFD) simulations are computationally demanding and therefore impractical in many cases. To enhance aircraft design, it is desirable to obtain models joining the best of these two worlds. A common approach is to distill high fidelity methods into Reduced-Order Models (ROMs) that can accurately approximate unsteady aerodynamics at orders of magnitudes lower costs than CFD. Relevant literature offers many different ROM techniques for varying purposes. Nonetheless, constructing such models is still challenging and currently there is no generally agreed method.

In the current thesis a ROM technique that may be applicable to wider ranges of problems and simpler to construct is sought. The objective is to obtain a model that can promote aircraft control design, performance assessment and structural analysis throughout dynamic maneuvers over complete flight envelopes. The thesis proposes a novel approach utilizing modern, deep convolutional neural networks (CNNs). The devised model consists of three main components. First it incorporates a geometry description constituted by coordinates of an aircraft CFD surface grid. Second, a primary encoding-decoding CNN predicts pressure distribution at the grid points of the geometry. The final and third part of the model is an auxiliary encoding CNN deriving integral aerodynamic loads corresponding to the pressure field predictions of the primary network. The model evaluates and produces instantaneous values. Given a maneuver, it proceeds in timesteps. The predictions of the separate instances are computed directly without the need of subiterations (as it would be the case for CFD simulations).

As a proof of concept, the model is applied to symmetric motions in the vertical plane at fixed Mach number and altitude. The subject of the investigations is the MULDICON configuration of the 251 th Science and Technology Organization work-group of NATO. To fully exploit the advantages of reduced-order modelling, flow characteristics are inferred from a single excitation following an efficient system identification technique using Schroeder sweeps as input signals. The performance of the model is assessed by numerous test cases performed in CFD. First, steady conditions of varying incidence angles are investigated. Second, harmonic pitch and plunge oscillations around different angles of attack at different amplitudes and frequencies are considered. Third, additional test cases of a linear pitch up-down -- and a climbing maneuver are studied.

Considering computational efficiency, the results show robust model performance. GPU-accelerated CNN calculations are conducted roughly 5000 times faster than CFD simulations. The primary network can accurately resolve the pressure distributions over large portions of the geometry. Lower surface predictions are very accurate. However, among certain conditions discrepancies are observable on the upper surface towards the wingtips. Still, the secondary network can predict corresponding aerodynamic forces accurately. In contrast, its moment predictions are sensitive to errors in pressure distributions. Consequently, moment predictions can largely deviate from reference data, especially when nonlinear phenomena are prominent. However, in many cases errors are attributed to insufficient regressor space coverage, i.e.\ certain input combinations are explored poorly by the Schroeder sweeps. Reconsidering system identification practices might mitigate those issues. Nevertheless, the thesis proves the applicability of deep CNNs to the problems at hand. Additionally, the results encourage further investigations.

In the year 2050, global anthropogenic radiative forcing from aircraft emissions are projected to increase significantly. Recent studies have considered climate optimized flight trajectories to be a promising measure to mitigate non-CO2 emissions’ environmental impact, which is highly sensitive to locus and time of emissions. Estimating the maximum mitigation potential from these trajectories requires accounting of air traffic regulations. As designing regulated climate optimal trajectories necessitates solving a hybrid optimal control system with unknown mode sequence and associated switching times, there is a need to build an efficient and systematic control technique. In this thesis, a bi-level optimal control algorithm is proposed for designing climate optimal cruise trajectories, the lower level calculates the optimal switching times and control
inputs of a fixed mode sequence and, the upper level updates the mode sequence with mode insertion which lower the cost locally. The problem for trajectory optimization is formulated here as a hybrid optimal control problem with a switched system and with a variable mode sequence, where step-climb and descent modes are included in the mode sequence. Optimal Control problems for minimizing operating cost and climate cost with fictitious climate cost functions (CCF), varying with altitude, are solved to study the performance of the algorithm. The algorithm is implemented within the Trajectory Optimization Module (TOM) by building a bi-level framework. The framework was validated by solving the operating cost optimal control problem. The maximum error between the cost reduction estimated by the algorithm and the actual cost reduction was found to be less than 15%. With high probability it can be stated that the bi-level framework is able to calculate an optimal mode sequence as the framework allow for zero entry modes in the mode sequence i.e. modes of zero duration. Although, careful consideration is required while selecting a mode for insertion as the framework is highly dependent on the sequence of the set of modes.

Despite a satisfactory performance of the bi-level optimal control technique there are few challenges which limits the scope of this technique. The maximum error was found to increase for optimal control problems with AirClim CCFs. The dependence of the AirClim CCFs on position of the aircraft influences the locus of the trajectory at each flight level. Because of this the the trajectories calculated in each iteration of the framework are found to be inconsistent. A flight trajectory guided by waypoints is proposed as a solution for future studies to handle the inconsistency between trajectories. As future studies are expected to focus on finding optimal mode definitions for designing climate optimal trajectories, the bi-level optimal control algorithm can act as an intermediary tool with which the researchers can systematically investigate cost benefits along the trajectories.

Deployable morphing aircraft are gaining popularity in the academic and industrial world. Their deployability offer the possibility to be stored and takes less space. This is advantageous in volume limited applications like space explorations, where every cubic centimeter is important. Another field where this aircraft is increasingly used is the military field. Deploying it in the battlefield directly from a carrier aircraft opens endless possibility to the operation of such aircraft. However, mid-air deployment is subjected to several internal and external factors that could lead to the failure of the deployment.

In this study, a method to assess the safety of deployment of a deployable morphing UAV is constructed. To do this, an aircraft test case is taken and a corresponding model of the aircraft is made using the Multibody dynamics approach. This model is then verified and validated by two different methods. First, a comparison of the stability derivatives of the created model and an off-the-shelf aerodynamic solver is performed. Secondly, the response of the created model is compared with actual flight test data, where the test case aircraft performs a maneuver. For both validation methods used in this study, the created model is able to resemble the output from the off-the-shelf aerodynamic solver and the flight test data. The validated model is then used as to test the method developed in this study.

The method developed starts with the definition of a safe deployment of a deployable morphing aircraft, where three different safety concerns are considered. The definition of a safe deployment also helps in deriving different categories of deployment which includes safe and various unsafe deployment depending on what causes the deployment to be unsafe. This method is tested on two different deployment scenarios.

By using the method developed in this study, safety deployment spaces are constructed from the range of input parameters determined for the two different scenarios. It is revealed from this safety deployment space what combination of input parameters is favorable for a safe deployment, and what causes the different unsafe deployments. Unsafe deployments are also categorized to understand which of the safety concern causes the deployment to be unsafe.

The flow field subjected to the influence of the helicopter rotor is characterized by its three-dimensional pattern and unsteadiness. The accurate modeling of the flow around the rotor blades addressed by Computational Fluid Dynamics (CFD) simulations is associated with high computational costs. The complexity of the analysis can be reduced by means of simplified methods such as the actuator disk approach, where the rotor is modeled as a zero-thickness porous surface.

The CFD solver TAU developed at the German Aerospace Center (DLR) includes among its features an actuator disk module which transmits to the flow field prescribed time-averaged load distributions both in axial and tangential directions by means of pressure and tangential velocity jumps. The computation of these loads can be achieved by using the Blade Element Analysis Tool (BEAT), a developed rotor code based on the blade element theory. In addition, since it is considered that changes in blade motion about the feathering bearing and flapping hinge induce variations in the aerodynamic loads acting on the blades and vice versa, the loads need to be calculated under trimmed or equilibrium conditions.

The coupling between BEAT and the TAU actuator disk is defined in a way that the velocity captured at the grid points of the actuator disk surface after each simulation performed in TAU is transferred to BEAT, which computes an updated aerodynamic load distribution. The performance of this approach is tested for an isolated rotor configuration in hovering and forward flight conditions. In hovering flight, the convergence of the CFD flow simulations towards the steady state solution is not satisfactory due to the stiffness of the compressible Navier-Stokes equations at low Mach numbers. Furthermore, reverse flow regions are determined by TAU at the inner and outer boundaries of the actuator disk. The recirculation flow entails high gradients in angle of attack between neighboring blade sectional elements, which yields to the unstable formation of new reverse flow regions. Nevertheless, in forward flight conditions the performance of the flow solver is robust and convergent solutions can be obtained. Moreover, as the flight speed is increased, the shed vorticity is displaced more quickly outside the rotor disk and, hence, its associated effects on the performance of the rotor are diminished.

The accuracy of the coupling approach is validated by comparing the computed results with those measured in a wind tunnel test campaign. The found differences in pitch control angles are assigned to the fact that the blade elastic deformations are neglected in the developed method. This statement constitutes the baseline to be developed for future work.

Finally, the reduction in computation time required by the coupling approach with respect to other more accurate methods enhances the idea of further developments. Therefore, the developed method can be regarded as a suitable strategy to tackle the problem in forward flight conditions in cases where high fidelity results are not needed, such as in the preliminary design stages.

The Hyperloop is a newly proposed transportation method which caters to the rising demand of high-speed transportation, while adhering to stringent emission requirements. This system comprises pods that are propelled through a partially evacuated tube at velocities of 300 m/s. The low-pressure environment paired with high-subsonic Mach numbers create an unconventional flow regime. This thesis project is an investigation on the aerodynamic characteristics of a Hyperloop pod, whereby an optimization is performed in order to find a pod shape that yields minimum drag. A low-fidelity solver is developed to perform the analysis and optimization. Investigations show a large pressure peak at the nose due to violation of the Kantrowitz limit. Furthermore, a blunt nose paired with a concavity in the tail section yield minimum drag due to the alleviation of the choked flow effect. The solver can readily be embodied into a larger, multi-disciplinary optimization framework.

During aircraft design,multiple tools are utilised to inspect the performance of the configuration. As the design matures, higher fidelity analyses are conducted to predict the flight dynamics of the aircraft. These analyses are conducted by using semi-empirical relations, numerically analysing flow behaviour, conducting wind-tunnel tests and performing scaled test flights. However, Semi-empirical relations might not hold for next generation aircraft and wind tunnel testing and scaled test flights are extensive and are also prone to accuracy issues. A best of both worlds can be found in numerical analysis. However, an increase in flow fidelity modelling comes with an increase in computational cost. Besides, complete analysis of all possible manoeuvres of a design increases the number of computations significantly. Current methods cope with this issue by using flight dynamics models based on so called stability derivatives, instantaneous values which couple flight state parameters to aerodynamic loads to predict aircraft flight dynamics. However, these models do not take into account time dependency. Therefore, these methods do not accurately predict the flight dynamics of agile aircraft, such as unmanned combat aerial vehicles, undergoing rapid manoeuvres where unsteadiness dominates flow behaviour. This conventional reduced-order modelling method, in which samples of the full-order model are taken in the form of stability derivatives, causes design iterations to be analysed inaccurately. The objective of this report is to investigate reduced-order modelling for flight dynamics prediction, thereby comparing conventional techniques to a method which does take into account unsteadiness in flow behaviour.
The method investigated is based on indicial step response functions, which are samples in the form of unsteady aerodynamic flow behaviour functions of the full-order model. The idea is that once these samples are known, any flight manoeuvre can be analysed within minutes. Research found in literature has assessed some of the capabilities and limitations of this method, but not yet applied this to flight dynamics prediction. The research described within this report will address this gap by using two test cases. The first testcase is used to assess the assumptions made in literature, on aerodynamics loads modelling, by applying the method on a two-dimensional airfoil in subsonic flow conditions. It was found that the indicial step response functions are indeed representing the full-order model, thereby taking into account unsteady flow behaviour in aerodynamic loads prediction. In longitudinal motions, the angle of attack and pitch rate effect need to be taken into account to predict lift, drag and pitching moments. Multiple frequencies of the same manoeuvre can be analysed within minutes once the samples are calculated. Results show that the accuracy of the predictions becomes a trade-off issue between samples calculated and accuracy required. The second testcase is used to apply the indicial response functions to flight dynamics prediction of an agile
unmanned bomber aircraft undergoing fast manoeuvres. A longitudinal-directional climbing manoeuvre was calculated by developing a flight dynamics model based on stability derivatives. The flow behaviour encountered during this manoeuvre was analysed to include highly unsteady and non-linear phenomena (e.g. vortices and flow separation) at higher angles of attack. By comparing the results of themethod under investigation to the full-order solutions, it was shown that aerodynamic flight dynamics predictions were accurate in capturing unsteady behaviour and weak non-linear flow behaviour. However, the samples proved to be inaccurate in representing behaviour in highly non-linear regions. Concluding, this means that indicial step response functions provide more accurate flight dynamics predictions than conventional stability derivatives in representing unsteady flow behaviour. The accuracy of the predictions are highly dependent on the samples chosen. Several samples suffice to predict the unsteady behaviour for linear and weak non-linear flow regions of the flightmanoeuvre. If surrogate modelling is applied, the method can become more computational efficient than conducting multiple full-order time-marching numerical calculations. It is recommended that more research is performed on indicial step response functions
in capturing highly non-linear flow behaviour, as the research showed that the size of the samples affects the flow behaviour representation.

Looking at performance prediction of small coaxial rotorcraft, it was found that there are no fast methods available for designers. This research was aimed at using a blade element method in combination with the momentum theory to quickly predict the performance at small coaxial rotorcraft. These methods are successfully applied in larger rotorcraft. It will therefore be of interest to see how these methods perform when the aerodynamics changes due to scaling. Especially the Reynolds number effects are expected to influence and potentially limit the application of the modelling methods.

The new developed model was tested on the full-scale Kamov 32 helicopter as well as two small Unmanned Aerial Vehicles. These are the commercial available Walkera Lama and the Guardian Angle, which was developed by undergraduate students during a design synthesis exercise. The results show a successful implementation of the calculation methods, with some interesting insight in the limits imposed by them.

Flight testing is an activity that exists since the beginning of flight, and is still seen as an essential part of the aircraft development process. While numerical simulation has been improved significantly over the last decade, flight testing remains attractive in the exploration and evaluation of the flight envelope, especially in the region where numerical simulation lacks prediction accuracy. Free-flying model testing is a potential technique, which is recognized by some of the largest research institutes. Using such relatively cheap scale models aids in the determination of the flight behavior of full-scale aircraft in an early stage in the design process. Furthermore it takes less time to develop a scalemodel and is also safer to fly. A 5.5% scale model of the Boeing 737-700 is created by the TU Delft as a first step towards investigating the potential of dynamically scaled flight testing. This Delft University Unconventional Concept (DUUC) has a conventional tube-wing configuration and features ducted fan engines mounted aft on the fuselage, which replace the conventional empennage. A horizontal and vertical jet vane is placed in the duct, which guide the propeller slipstream and thereby control the attitude of the aircraft. Successful flights have already been performed with this model, which proves the propulsive empennage concept. However, the stability and control behavior has only been assessed at a low level. Flow characteristics entailed by the ducted fan are accounted for by means of approximate functions and empirical relations instead of looking at the physical interaction between the components within the duct. This study focusses on the processing of said interaction by performing a CFD study of which the results are used in the longitudinal equations of motion of the DUUC. This way, certain motion can be simulated from which the effect of the unconventional empennage on the stability and control can be assessed. In order to study the aerodynamics of the propulsive empennage and subsequently its stability, a CADmodel is developed using PARAPY. This is a knowledge based engineering platform written in the Python language and allows for parametrization of the model such that geometric alterations can easily be implemented. Previous studies within the TU Delft have lead to a multi-model generator written in PARAPY. The fuselage and wing generation modules that were developed can readily be applied in the generation of the DUUC model. The propulsive empennage module is created by the author of this thesis, consisting of several sub-parts: (1) outer duct that forms a shielding around the propeller, (2) the pylon that connects the duct to the fuselage, (3) a center body to which the (4) propeller and (5) control vanes are attached. The full aircraft model has been used in a series of CFD simulations with ANSYS FLUENT. Inviscid steady calculations are performed using a pressure-based solver with a SIMPLE pressure-velocity coupling scheme. The propeller is modeled as an actuator disk whereby the thrust is defined by specifying a pressure jump across the thin surface. The thrust produced by the ducted fan is based on the turboprop installed on the ATR72-600 aircraft, which has similar dimensions compared to the Boeing 737-700. Both the full-scaleDUUC as well as the 5.5% scale model are subjected to this aerodynamic analysis. Minor discrepancies were found between the two models, which is a result of the artificial viscosity that is implemented in the Euler scheme to smoothen strong gradients in the solution (numerical dissipation). Due to the location of the engine, the thrust line is located above the center of gravity which contributes to a stabilizing effect. The effect of several parameters on the stability and control has been analysed, such as the center of gravity location, mass and inertia, and geometric scaling factor. The slight difference in aerodynamic performance between the full-scale and subscalemodels, means that the trimmed solution for steady horizontal flight is also slightly different with a maximum error of 2.5 degrees in elevator deflection. More drag is obtained as the aircraft scale goes down, resulting in a higher required thrust to achieve force balance in a trimmed flight condition. Furthermore it is found that by scaling the aircraft down according to Froude scaling theory, the response becomes much faster. This response can in turn be scaled back, which proved to be a very useful estimate of the full-scale simulation.

The solution space for conflicts between aircraft consists of velocity, heading and altitude changes. In the current design of the air traffic controller workspace the instantaneous values for velocity, heading and altitude can all be found in the aircraft labels on the radar, however the heading and velocity can also be derived from the direction and speed of the aircraft on the radar display. Since the radar is a top view of the traffic situation, no means of deriving the altitude other than the aircraft labels can be found. Previous attempts to present the altitude domain in the radar display showed room for improvement. These display designs have been analysed and information gathered from them is used to create a new ecological user interface. The aim of this interface is to improve the support in conflict detection and resolution for air traffic controllers by representing the situation in the altitude domain. A preliminary feasibility test was performed from which it can be concluded that the altitude display has potential to improve safety in air traffic control and to reduce the complexity within the sector, but that some more attention should be put on making sure the aircraft leave the sector within the constraints linked to their exit waypoint. The subjective data from the questionnaire at the end of the feasibility test showed that the users felt that generally the display was easy to use.

Aircraft design methodologies have been significantly developing from the past few years with the advancements in knowledge based techniques. These methods enable the storage of design knowledge and rules, and reuse them to create different types of designs, thus preventing the designer to perform repetitive tasks. Tasks such as parametric modelling of components, such as the aircraft wing can be automated by storing the modelling processes and the design rules in a knowledge base. With this process, variants of the wing with different geometric parameters can then be generated in a short duration by simply varying certain top-level requirements. It is necessary to extend these design techniques to model aircraft systems in the conceptual design stage. This, not only decreases the time of design realisation but also presents a scope to assess the effects of various inter-dependencies due to systems and make appropriate changes, in the early stages of aircraft design. Developing and demonstrating a framework which aids to assess the influence of the wing subsystems, namely the flight control actuators, fuel tanks and anti-ice elements; on the aircraft design and performance in the conceptual design stage is the aim of the thesis.

This thesis presents a combination of physics based and knowledge based design methodologies to size the wing subsystems and position them in the airframe. Consequently, the methods are integrated into the conceptual aircraft design process to enable multidisciplinary design with supporting domains. The methods are aimed to aid the design of conventional systems architectures and More Electric Aircraft (MEA) systems architectures as well. With these methodologies, the Systems Model Generator (SMG) application is developed in Python to facilitate semi-automatic wing subsystems sizing and orientation in the airframe based on top-level aircraft requirements, initial aircraft design parameters and system specific parameters. The subsystem models generated with the proposed methodology for short-medium range civil transport aircraft are verified and validated as well. Knowledge based systems and subsystems selection are implemented to facilitate semi-automated systems, subsystems and architecture selection, based on the aircraft configuration and systems specific requirements. Methods for automatic iterative fuel tanks sizing and intersection detection are implemented to further reduce the overall design time and make the tool more suitable for integrated sizing.

With the multidisciplinary design framework, the conceptual parametric models, volume, mass, power consumption and position of the subsystems in the airframe are generated and propagated in the conceptual aircraft design stage; thus bridging the conceptual and the preliminary design stages. In the proposed framework, the domains of aircraft design generation, systems selection and sizing, subsystems selection and sizing, engine sizing and mission simulation are considered for the multidisciplinary design process. The domains are integrated with the DLR CPACS-RCE framework.

A case study to demonstrate the process of integrated parametric subsystems sizing of the aircraft, with the proposed framework is presented. The aim of this case study is to assess the influence of the MEA systems architecture relative to the conventional systems architecture for a short-medium range transport aircraft, similar to the Airbus A320-200. In this case study, the quantitative influence of the subsystems' parameters on the aircraft design and performance parameters is determined and analysed. The subsystems' parameters constitute the mass, power consumption, volume and location of the subsystems in the airframe and the aircraft design parameters constitute the aircraft masses such as the overall empty mass and the fuel mass for the mission. The generation and propagation of the design and performance parameters of the aircraft through each domain of the framework are presented and analysed as well with the case study. In this case study, it is observed that the MEA systems architecture results in a lower mission fuel mass relative to the conventional systems architecture by nearly 2.3\%. Furthermore, these results are compared with literature and observed to be in the similar range of 2-7\%. Thus, the validated aircraft design framework presented in this thesis enables to substantially increases and propagate the design knowledge of aircraft systems, in the early design stages.

Helicopters are complex and expensive aircraft with a level of technology that is immature compared to fixed- wing counterparts. Helicopters suffer from vibratory loads stemming from the main rotor and exhibit control and stability problems in the low-speed flight regime. Operating a helicopter near the limits of its flight envelope may result in unacceptable high structural loads which adversely affects wear and tear of drive train components. Conservative safety regulations lead to high operating cost where a significant part can be attributed to maintenance. Manufacturers and operators therefore strive to make helicopters more capable and reliable in an effort to reduce operating cost. Structural load alleviation offers an attractive option to achieve this goal by reducing component damage accumulation and subsequent required maintenance.

This thesis investigates structural load alleviation in the tail rotor drive train of the UH-60 Black Hawk. The Black Hawk provides a compelling case for load alleviation research because of its ever growing operational weight and resulting increase of drive train load levels. Furthermore, the lifetime of the UH-60 is to be extended so that it will fly for many years to come. Current research and applications of rotorcraft structural load alleviation focus on the main rotor but less attention is given to tail rotor drive train components. This project seeks to address this knowledge gap by investigating manoeuvres that result in critical dynamic loads in the UH-60 tail rotor drive train. A survey of pertinent literature and interviews with helicopter pilots indicate that pedal inputs for left-hand turns in hover lead to high dynamic loads in the UH-60 tail rotor drive train.

A flight simulation model is constructed that offers the novel capability to predict dynamic loads in tail rotor drive shafts. This model consists of an available high fidelity engine model and existing rotor models coupled by a multi body dynamics tail rotor drive train model with properties that are based on component measurements and CAD drawings. Experiments are conducted to determine the relation between manoeuvre aggressiveness and dynamic loads in tail rotor drive shafts. Based on the results a manoeuvre load alleviation control strategy is devised to reduce dynamic loads while ensuring applicable Level 1 handling quality requirements. Application of this control strategy will decrease dynamic loads during left-hand yaw manoeuvres in hover. Furthermore, the results highlight what reduction in loads can be achieved for varying levels of manoeuvre aggressiveness. These findings may aid in the design of flight control systems that incorporate tail rotor drive train load alleviation objectives.

The objective of this research is to determine the impact of design variables on power output of the Makani Power Wing7 airborne wind turbine. This system consists of a rigid wing connected to the ground using a using a number of bridles and a tether. The tethered wing flies crosswind trajectories and generates power using on-board turbines. Sensitivities of power output to design variables are required for various design trade-offs. A simulation and optimization approach is presented which allows analysis of a fairly complex model with realistic controls and constraints. Using trajectory optimization removes the assumption of a predefined trajectory and allows analysis of the impact of the flightpath itself. The wing is modeled as a rigid body with six degrees of freedom. Aerodynamic forces are determined using a response surface on results of a vortex lattice code. A time-varying position of the bridle attachment point couples the wing to the tether. This tether is implemented as a straight elastic element with analytic approximations for the effects of mass and aerodynamic drag. The power system model consists of a turbine response surface and DC-equivalent generator model. Optimal trajectories are determined using the direct collocation method resulting in a capacity factor of 48.1% for a site with uniform wind field and annual mean wind speed of 7.5 m/s. The relative impact of aerodynamic lift and drag on long term averaged power output are found to be an order magnitude more important than wing mass. Topology and circularity of the flight path are found to have a low impact on power.