Considerations that decide producibility of a design are an important part of the design process, and must be included in the early design stages to ensure that these designs can be realised. Not including these considerations carries the risk of incurring additional costs and delays at later stages of product development because of design changes, or can lead to limiting oneself to conservative design choices to reduce the associated risk.
The current process in the industry accounts for these production considerations in design through a manual process that is iterative and time-consuming, and hence forms a bottleneck in being able to trade-off multiple design concepts. Attempts at accounting for these production considerations in an automated way are associated with the limitations of either only considering the manufacturing cost, being specific solutions that work only in certain scenarios, or being dependent on some commercial software tools, which are not fully suitable for use in context of automation and/or at the conceptual design stage. Additionally, the aspects of manufacturing and assembly are usually not considered at the same time in these studies.
Therefore, this thesis aims at developing a methodology that enables the automated inclusion of production considerations in the conceptual design process of aircraft structures, while overcoming shortcomings of the state-of-the-art....

Interest grows rapidly in electric and hybrid electric aircraft. To determine the optimal performance and energy management required with such novel powertrain configurations, a knowledge-based aircraft and powertrain performance model is developed. The model is then used to set up an optimal control problem, which is transcribed to a non-linear programming problem using global orthogonal Legendre-Gauss-Radau collocation for single phase problems, and Hermite-Simpson local collocation for multiphase problems. The solution to the control problem allows identification of the best control strategies and energy management strategies. A case study is performed on the HY4 hybrid fuel cell aircraft and the hybrid electric Pipistrel Panthera. Solutions show that for best fuel economy, flying at a minimum drag airspeed, and keeping a constant power setting, proved more important than the choice of altitude. This was more noticeable for the HY4, with its relatively low power available and good aerodynamic properties following from its glider-based airframe.
The Fuel-optimal energy management strategies proved identical for both aircraft investigated. Batteries are used to provide a power boost during takeoff, after which batteries are discharged gradually throughout the remainder of the flight to maximize discharge efficiency. The engine or fuel cell are kept at approximately constant cruise power settings throughout the flight. The Panthera showed
consistent flight profiles with increasing range. For the HY4, however, achieved airspeeds reduced with increasing range, and additional measures were required to force a climb to non-zero altitudes due to its under-powered nature. The fuel-optimal trajectories offered an average of 10-15% of possible fuel
savings, depending mostly on the size of the onboard batteries. Fuel savings increased significantly at low ranges300km, where the contributions of the batteries have more impact.
Comparing different transcription methods and problem setups, it was concluded that global orthogonal, or pseudo-spectral, methods like Legendre-gauss-Radau collocation are not only faster, but also more consistent compared to simpler direct collocation methods. However, if the problem complexity increases and the performance limits of the aircraft are pushed, switching to a simpler method like Hermite-Simpson collocation reduced the time required to find a solution, with negligible differences in the resulting trajectories. Opting for a multiphase problem set-up, essentially splitting the problem in a series of individual subproblems, appeared less advantageous. While offering more control over the trajectories, time required to find solutions increased drastically, and offered no additional insight into the best energy management strategies.

Fuel cells are one of the promising technical solutions for the first generation of sustainable aircraft. Currently, there are several shortcomings in the literature related to the conceptual and preliminary sizing of the propulsion systems of such aircraft: not only is there a lack of general methods, but also, much less effort is put into the estimation of the system’s weight, as compared to its performance. In this thesis project, a set of preliminary sizing methods was developed for the propulsion systems of CS-25 proton-exchange membrane (PEM) FC aircraft. These methods were then integrated into the Initiator, an aircraft sizing environment. With this approach, a complete preliminary sizing procedure of a regional turboprop aircraft was successfully performed. Finally, sensitivity studies were carried out in order to demonstrate the advantages of this modelling approach.

Classically, aircraft controls are designed such that every control surface type primarily influences a single degree of freedom by creating a moment. Increased availability of computational resources and novel aircraft configuration allow a deviation from this approach and to utilize individual control surfaces to generate moments around multiple axes. The research investigates the impact of control allocation algorithms on the required control surface span and area for a box-wing configuration aircraft, the PrandtlPlane. The unique geometry of two full wings allows more flexibility in control surface placement. An optimization system for automatic control surface sizing under the constraints of adequate handling qualities has been developed and used to compare mechanical gearing, the constrained pseudo inverse, and the direct allocation algorithms. The results show that the PrandtlPlane configuration can benefit from the use of control allocation, showing a clear advantage of the direct allocation algorithm.

The demand for air travel is increasing as more people gain access to commercial aviation. As the current generation of aircraft makes use of fossil fuel combustion, this growth results in an increase in the global emissions. This is at odds with the worldwide efforts of reducing the adverse effects of climate change. Therefore, advanced propulsion systems must be developed to limit the emissions caused by the commercial aerospace industry. Using hydrogen fuel cells for propulsion is a promising technology to potentially get the sector to zero emissions. It is of interest to explore the capabilities and feasibility of aircraft with a hydrogen fuel cell powertrain. To determine the feasibility for a wide range of aircraft, a general design methodology is required.Current research efforts focus on component level performance, however system level design research, while present, didn't introduce a general methodology. The most pressing challenge was found to be related to the system level design of a CS-23 category hydrogen fuel cell aircraft. The CS-23 category aircraft class has been identified as the most suitable focus for research efforts, due to the lower technical and certification requirements placed on the components to reach a feasible design. A general methodology for the design of CS-23 category aircraft was therefore found to be a useful contribution to the state of the art. This additionally provides a deeper understanding into the most important parameters of the power and propulsion systems design.In this report, a general methodology for the conceptual design of hydrogen fuel cell powered CS-23 category aircraft is presented. The methodology makes use of a modified class 1 weight estimation for initial sizing according to customer requirements and component technology levels. The generated aircraft is refined using further aerodynamic analysis, which results in a feasible aircraft concept, as well as component level specifications for important aircraft components.The methodology is implemented in a software tool, HAPPIE (Hydrogen Aircraft Power \& Propulsion Initial Estimator), which allows for rapid sizing of different hydrogen fuel cell concepts. This makes the methodology accessible, and additionally provides feedback on the effects of individual design choices and technology levels on system level performance. SUAVE is used to perform the refined aerodynamic analysis.The methodology is validated using existing conventionally powered aircraft, by comparing the sizing results from the methodology with publicly available data. The methodology's ability to analyse conventional as well as hydrogen fuel cell powertrains, furthermore allows for performance comparison between current and future technologies.The results of the sizing methodology demonstrate the viability of hydrogen fuel cell aircraft in the CS-23 category. A conceptual design is generated, which serves as a baseline for the sensitivity analyses. It is found that hydrogen fuel cell aircraft are generally heavier than conventional aircraft, using current technology levels. Liquid hydrogen is identified as the best hydrogen storage method. Compressed hydrogen storage is also possible, however this results in a heavier aircraft with limited range. The current methodology does not predict thermal behaviour to have a significant effect on the mass of the aircraft. A component sensitivity analysis determined that the fuel cell efficiency, fuel cell specific power and hydrogen storage efficiency are the most important parameters.The ideal cruising altitude for fuel cell aircraft is at an intermediate altitude, due to the fact that the fuel cell powertrain performance decreases at increasing altitude, which balances with the lower drag in lower density air.The research demonstrates that hydrogen powered CS-23 aircraft are viable for current technology levels, and are a suitable way in reducing carbon emissions in this category.

The medium-altitude long-endurance (MALE) unmanned aerial vehicle (UAV) plays a key role in both military and civilian applications around the world. However, in a world where electric UAVs are increasingly becoming the norm, the MALE UAV is still primarily powered by hydrocarbon fuels and its design is very much constrained by existing general aviation reciprocating engines. The aim of this study is to investigate the design effects resulting from incorporating electric propulsion systems in MALE UAVs, while also investigating the effects of modelling fidelity on the results. Reciprocating engine, hydrogen fuel cell and battery models are incorporated into a multidisciplinary design optimisation framework, with the aim of minimising the maximum take-off weight at various levels of modelling fidelity. The results show that hydrogen fuel cells can provide around 40% reduction in the aircraft's maximum take-off weight, while batteries are not capable of enabling long-endurance flight. It is further shown that the effects of modelling fidelity of the aerodynamics and structural wing weight models are more significant when the aircraft design being optimised is substantially different from the baseline configuration.

It is the aim of this research to assess the mission performance of a boxwing aircraft by developing a configurationagnostic, multifidelity optimal control toolbox for performance and mission analysis. The boxwing aircraft, sometimes named a PrandtlPlane (PrP), is an unconventional aircraft. An instance with redundant controls is designed within the PARSIFAL project. This specific aircraft is designed for commercial transport in the shortrange segment (a 4000 km design range) and for a high passenger capacity (up to 308 passengers). Because of the beneficial induced drag characteristics inherent to the boxwing configuration, the PrP represents a possible solution towards the sustainable future of aviation. To investigate the potential of the PrP as an alternative to conventional commercial aircraft, the mission performance assessment of the aircraft has been split into two components. The first part of the assessment covers the comparison between the performance of the PARSIFALdesigned PrP and that of a competitor aircraft with a similar design range, the A320, while allowing only nonredundant controls. The second part of the assessment involves the quantification of the PrP’s performance when allowing redundant controls in the form of Direct Lift Control (DLC), enabling the aircraft to increase its net lift without a change in pitching moment. Analyses of the PrP and its competitor aircraft for various ranges have shown that the PrP outperforms its competitor in terms of relative fuel consumption. When flying its minimumfuel mission, the PrP’s competitor consumes less fuel in absolute terms. Nonetheless, because the PrP carries more than twice as many passengers, it consumes up to 14.5 % less fuel per passenger per kilometre. In other respects the PrP’s performance is inferior to that of its competitor. The 5400 km maximum range of the PrP is considerably lower than its competitor’s maximum range of 6200 km. Moreover, at a fueloptimal Mach number of approximately 0.7 the PrP cruises appreciably slower than the cruise Mach number for which it was designed, unlike its competitor. In general, the PrP flies its trajectories much slower than its competitor at an approximately 10 % lower average velocity in the minimumfuel missions. If both time and fuel are considered equally in the cruise altitude optimisation, the design altitude of 11 km is deemed appropriate. If only fuel consumption is considered, the PrP would benefit in fuel economy from lowering the initial cruise altitude at the cost of increased mission time. At an optimal altitude of 9.3 km, the PrP would consume 2.2 % less fuel than at its design altitude of 11 km at the cost of even slower flight. The sensitivities of the PrP’s mission time and fuel performance to changes in its design Zerofuel Mass (ZFM) have been investigated. Keeping the Maximum Takeoff Mass (MTOM) constant while varying the ZFM, design mission simulations were run for the PrP for several objective functions. It was found that when flying for minimum fuel, a 1 % increase in ZFM incurs a fuel consumption penalty of over 1 % through a nearlinear, direct proportionality. Likewise, the mission time varies nearly linearly with the ZFM; a 1 % increase results in an approximate mission time increase of nearly 0.5 %. The incremental aerodynamic lift and drag due to control surface deflections for DLC were modelled using a flatplate approximation. With this approach, the projected missionlevel benefits of using DLC are marginal. On the design mission, the results indicate an increase in fuel economy of 0.6 % on the minimumfuel mission and negligible temporal gains on the minimumtime mission. It is however emphasised that numerical uncertainties due to the discretisation of the problem pollute all obtained solutions to some degree, such that appropriate caution should be exercised when interpreting these results in an absolute sense. In future research, a grid refinement study would be a valuable addition to quantify and bound these uncertainties. It is deemed equally important to look into a more sophisticated way to model the control surface aerodynamics necessary for assessing the benefits of DLC. A broader recommendation pertains to future research on boxwing aircraft aerodynamic design. The current research has indicated that the optimal trajectories for the PrP result in very distinct flight profiles when optimising for different objectives. Therefore, it would be interesting to see how the aerodynamic design could evolve, such that flying for fuel economy wouldn’t require such a compromise in temporal performance and vice versa.

This research is about the integration of hybrid and electric aircraft (HEA) in the air
traffic management system. Current conventional aircraft are responsible for emissions and noise that often lead to nuisance for residents. In order to make aviation more sustainable, aircraft manufacturers are studying the possibilities to replace fuel with a battery. However, due to a different powertrain, the performances deviate from conventional aircraft. This affects the air traffic management operations. This research studies how to integrate hybrid and electric aircraft in between conventional aircraft at (crowded) airports.

With the global increase in passenger traffic and growing popularity of long-haul routes over the Asia Pacific region and Atlantic Ocean, the possibility for hypersonic transport could become an attractive option to reduce flight time over long distance from 16-20 hours down to around 4-5 hours. In this thesis, a Multi-Disciplinary Optimisation platform has been developed to allow for the optimal sizing of hypersonic transport vehicles using vehicle take-off mass as the performance indicator subjected to fuel volume and payload height constrains. The current platform is applied to the LAPCAT A2 hypersonic long-range transport configuration by Reaction Engines, to determine the impact of range and cruise Mach number on the design of hypersonic aircraft. Results show that the optimal shape is greatly dependent on the aircraft range and fuel volume constraint. Additionally, the optimum hypersonic cruise Mach number is dictated by a trade-off between mission time, engine efficiency and Thermal Protection System mass.

The aviation sector and its associated activities contribute to climate change, damage to the environment, community noise, and local air pollution. As a consequence it affects the health and quality of life of citizens that live close to airports, and through climate change every person on earth. To reduce the impact of air travel and thereby accelerate growth, new technologies are being investigated. One area of research looks into the opportunities of using batteries and electric motors as additional energy source and power converter. Besides solving some of the current problems, these new technologies could enable radical new forms of aviation and create new business models such as on-demand personalized air travel.

Because the design rules of novel hybrid-electric aircraft differ significantly compared to that of conventional aircraft, new methods must be developed tailored for this new technology. Due to the additional variables introduced by using hybrid-electric propulsion systems, a simple sizing study quickly becomes a multi-disciplinary optimization problem. Furthermore, these new propulsion systems should not be benchmarked in an isolated way due to synergistic benefits with other disciplines. To close the gap between clean sheet designs and retrofits, it has been concluded that research focusing on feasible concepts in the near future is required.

The objective of this research is to define the implications of using hybrid electric propulsion systems in general aviation aircraft. The influence of design, operational and mission choices on optimal performance are identified through optimization studies. The design space includes the geometry of the main wing and the operational variables of the propulsion system. The performance is either measured in terms of fuel or total energy consumption per kilometer for design ranges between 50 and 1000 km.

To perform the quantitative analyses, a conceptual design framework has been developed. Models of all propulsion systems have been developed that simulate their performance. Every part of the tool is either verified with experimental data or has already been verified in literature. A quasi-three-dimensional aerodynamic solver has been modified to determine the aerodynamic characteristics with minimal computational time. The battery performance is modeled as function of time and rapidly sized such that the required power can be delivered by the smallest battery possible. A multi-disciplinary optimization approach has been used to integrate all modules and converge to an optimal design as quickly as possible.

First of all, a difference is found in terms of optimal configuration and operation of an aircraft when designed specifically for minimal fuel or for minimal total energy, leading to a maximum difference of 5% in terms of energy consumption. A trade-off between designing for aerodynamic efficiency versus the ability to carry batteries is part of the optimization routine. It has been shown that aircraft optimized for fuel consumption carry systematically more batteries compared to aircraft optimized for total energy consumption, always leading to a heavier aircraft with larger wings. Regardless of objective, the aerodynamic efficiency becomes more important for an increasing range while the ability to carry batteries decreases.

Three variables have been defined that determine the amount of hybridization in terms of power and energy without the need for constraint functions, as any combination of these variables lead an inherently feasible design: a climb coefficient that determines the rate of climb as a fraction of the maximum rate of climb, the fraction of the cruise that needs to be performed full electric, and a hybridization factor that determines the throttle of the internal combustion engine.

Given a certain required power at the power management system, there exist one specific throttle that leads to the maximum effective efficiency of the internal combustion engine. This throttle depends on the round trip efficiency of the battery charge-discharge cycle. When the internal combustion engine is scalable, the optimizer always makes sure that the most efficient throttle matches the power requirements during cruise, i.e. it avoids recharging.

The maximum possible range is highly determined by the required cruise velocity, given a fixed internal combustion engine. By scaling the internal combustion engine such that it is powerful enough to deliver the required power for cruise, the maximum range can be significantly increased. Furthermore, the optimal cruise altitude is found to be increasing with range and cruise velocity. At a cruise speed of 60 m/s the optimal cruise altitude coincides with the lower bound of 2 km whereas the optimal altitude at a cruise speed of 90 m/s increases from 2 km to 4 km for design ranges from 350 km to 700 km. In the latter case, the cruise altitude settles at 4 km for longer ranges.

The maximum range at which full electric cruise is possible is found to be the most efficient range in terms of energy consumption. This range increases linearly with the battery specific energy, providing designers a good initial estimation tool.

This research adds to the scientific body of knowledge as it presents a method that solves the multi-disciplinary optimization problem associated with the design of hybrid electric aircraft. Furthermore, through multiple optimization studies it provides insight in the influence of the design choices, operational choices, constraints, and mission profiles on the optimal performance of general aviation hybrid electric aircraft.

One of the novel configurations that could revolutionize the aviation industry is the PrandtlPlane. The closed-wing design features a low front wing and a high rear wing, connected by a lateral surface. In this research, the rear wings of the configuration are attached to a set of fins. Previous conceptual closed-wing design studies have failed to accurately predict the wing weight of the configuration, by using empirical relations or semi-analytical methods designed for conventional wings. These methods fail to capture the three major structural characteristics of a closed-wing design: an over-constrained structure, significant secondary bending moments and shear forces, and a different lay-out of the constraints definition. Therefore, a semi-analytical closed-wing weight estimator was developed. The primary weight is estimated analytically, while the secondary weights are defined by empirical relations. The primary structure is sized using an equivalent beam method, which is designed to withstand the aerodynamic, inertial and fuel loads applied to the structure. The aerodynamic loads are estimated with a vortex lattice method, at a 2.5g pull-up manoeuvre. For the PrandtlPlane configuration, the fin is modelled as a support to the rear wing. Due to the over-constrained nature of the structure, a stiffness iteration loop is implemented in which the internal loads are determined using the displacement method and all cross-sections are sized. The cross-sectional design features four booms, which are cross-coupled and four skins, which are sized independently. A total weight estimation of a closed-wing design takes 20-30 seconds. Sensitivity analyses are performed to establish the main drivers of the total wing weight and its distribution. Wingspan, wing sweep angle, fin sweep angle and longitudinal center of gravity position were determined to have the most influence on the total wing weight and its distribution. The new tool was implemented in the conceptual design workflow of the Initiator. A comparison between a design study in the Initiator with the closed-wing weight estimator and the previous methodology showed large discrepancies. Three design studies were performed with missions varying in payload weight, number of passengers and range. The total offset in the wing weight between the two methods ranged from 4.5\% to 30.4\%, leading to discrepancies in the required fuel mass of 1.6% and 5.7% respectively. Apart from the large offsets in total wing weight, the previous methodology also failed to accurately predict the distribution of the weight between the front, rear and lateral wing. Furthermore, the offsets between the methodologies are case-dependent and no clear relationship between them can be distinguished. Future conceptual design studies should thus include a closed-wing weight estimator. Finally, a parametric study was performed to identify the effect of altering the wing area ratio between the front and rear wing on the required fuel mass. For a single aisle, medium range mission, a design with and area ratio of 1.25 is 2\% more fuel efficient compared to a design with an even area distribution between the front and rear wing.