The objective of this thesis was to further the validation effort on the SU2 flow solver for non-ideal compressible flows. To do so, an uncertainty quantification infrastructure based on the V&V20 validation standard was developed. To determine whether a model error was present, the different sources of uncertainty were combined and then compared to the error between the SU2 simulations and the experiments. A variety of paradigmatic test cases were proposed which, together, would constitute a robust validation of the SU2 flow solver. Two paradigmatic unit test cases were analyzed; namely, a supersonic inviscid flow and a Mach 2.0 flow over a 2.5 degree wedge. The operating conditions of the experiments were as follows: an inlet total temperature and pressure of T_0=252°C and P_0=18.4 bara, and a back pressure of P_out=2.1 bara, for the first case, and T_0=180°C, P_0=3.5 bara and P_out=1.0 bara, for the second. The solver is able to predict the shock wave angle accurately. This step towards the validation was successful because the calculated total expanded uncertainty of 1.79% is higher the comparison error of 0.57%.In the second part of the thesis, the uncertainty quantification infrastructure was adapted to a test case constituted by more complicated flow features; namely, a supersonic flow through a 5-channel blade row. The computed uncertainties were on the same order of magnitude as in the supersonic inviscid flow test case. The simulations were verified with state-of-the-art CFD software for the quantities pressure and the Mach number. The positive result of the validation represents a step forward in the validation of the flow solver for non-ideal turbomachinery cases. The adaptation of the infrastructure to the cascade blade row paves the way for complex cases and for system response quantities of interest in industrial applications of CFD software, such as efficiency, to be assessed.

Flow unsteadiness caused by impeller rotation, vortex-shedding, secondary flows etc. can lead to the generation of acoustic waves within the turbomachinery cascade. This causes pressure loading on the impeller. When acoustic resonance occurs, i.e. the frequency of acoustic wave excitation matches with the structural natural frequencies of the impeller, high fatigue and vibrations are encountered, which can lead to structural failure. Centrifugal compressor applications like turbocharging and process engineering require an advanced understanding of the aeroacoustic excitation mechanisms as these have been suspected of playing a significant role in structural failures. Given that the current state-of-the-art analysis methods are incapable of explaining various instances of structural failures, a novel Lattice Boltzmann Method (LBM) based approach is explored. For the first time, aerodynamic and performance predictions, along with aeroacoustic amplitudes from the LBM based approach will be compared with a conventional Unsteady Reynolds Averaged Navier Stokes (URANS) based approach as well as test rig data. This will assess the feasibility of the LBM model for analysing forced response behaviour of a centrifugal compressor operating at conditions of acoustic resonance. The research compressor has been chosen based on an aeromechanic test campaign where high impeller blade trailing edge vibrations were measured. The computational domain consists of full compressor wheel including the hub and the shroud cavities. An attempt will be made to quantify the resonant amplification factor by simulating off-resonant conditions. The findings from this research would result in the development of a numerical framework for assessing the physics and severity of resonant excitations in centrifugal compressors. It will also highlight the importance of accounting for aeroacoustic mechanisms in the aeromechanical design of centrifugal compressor stages.

In the pursuit of reducing the climate impact of aviation, there has been an increased interest in the adoption of renewable energy technologies. Examples of revolutionary technologies include hydrogen fuel cell systems and waste heat recovery using the organic Rankine cycle. The design of viable energy conversion systems for aviation poses unique challenges in terms of efficiency, weight and size. To this end, research on small-scale turbomachinery operating at high rotational speeds is increasingly pursued in the context of, for example, fuel cell air compressors or organic Rankine cycle turbines. Such machines typically call for oil-free operation to avoid contamination of the process fluid. Gas foil bearings can prove to be an enabling technology due to their reliability, oil-free operation and compatibility with high rotational speeds. The use of gas foil bearings to support organic Rankine cycle turbines requires lubrication with complex working fluids at operating conditions near the saturated vapour line or thermodynamic critical point. Although there has been an increased interest in high-pressure gas lubrication in recent scientific literature, there is still a lack of understanding of the effects of non-ideal compressible flows on the performance of (gas foil) journal bearings. In order to further address this knowledge gap, this work focuses on the modelling of such bearings lubricated with dense fluids made by complex molecules like those adopted for waste heat recovery at high temperatures in aviation. The compressible Reynolds equation governing the thin-film flow within the bearing is discretized using a finite difference method. The solution to the non-linear problem is obtained using a relaxation method in which a thermodynamic software program updates the non-ideal thermodynamic state properties after each iteration. The load-carrying capacity of the bearing is obtained by integrating the resulting steady-state pressure field around the rotor shaft. A perturbation method is applied to obtain the stiffness and damping coefficients used in a linearized rotor-dynamic analysis. The developed computational tool allows for the analysis of bearing performance under varying operating conditions. The conclusions of this work emphasize the challenge of generating sufficient load-carrying capacity and rotor-dynamic stability associated with high-pressure gas lubrication in journal bearings. Bearings operating with compressible lubricants near the thermodynamic critical point are typically characterized by turbulent thin-film flows with non-negligible molecular interactions. Reduced peak pressures within the gas film are anticipated, resulting in a reduced load-carrying capacity as compared to ideal gas lubrication flows. It is shown that non-ideal thermodynamic effects have an impact on rotor-dynamic stability by affecting the steady-state attitude angle of the bearing. The modelling of a gas foil journal bearing used to potentially support the turbine of the organic Rankine cycle hybrid integrated device (ORCHID) of the TU Delft has finally been considered. The results show the utility of a numerical model in assessing bearing performance and understanding the associated physics. This work can be used as a basis for future analysis and design of gas foil journal bearings lubricated with high-pressure process fluids.

The growing environmental concerns has put a lot of political & ethical pressure on the conventional industrial practices. The research community is ever so eager to investigate new technologies that can offer an efficient and environmentally sound solution to the issues caused by the polluting industrial processes. One of the potential solution is an Organic Rankine Cycle (ORC) which is proven to be an effective technology to efficiently extract energy from low-temperature sources. The ORC is basically a Rankine cycle but employs a high molecular weight organic fluid as the energy carrier. The peculiar non-classical gas dynamics behaviour of these fluids during dry expansion close to the vapour saturation line (in the dense gas region) poses certain challenges on the turbine design. Furthermore, the low speed of sound in organic fluids and high expansion ratio required in an ORC turbine leads to highly supersonic flows in the turbine. Such high speed flows within a turbine stage are susceptible to high shock losses, if necessary corrective steps are not taken during early design phase. This has sparked a wave of interest in the scientific community to come up with an efficient supersonic ORC turbine design guidelines. The design of a supersonic turbine rotor is crucial to ensure efficient performance of the turbine. In the mid-20th century, a supersonic impulse rotor design methodology for axial turbines was proposed that boasts shock-free turning of supersonic flows using the vortex-flow theory. This procedure relies on Method of Characteristics (MOC) to solve hyperbolic governing PDEs. The method got enriched later with the capability to account for dense gas effects on the rotor geometry which is essential for ORC applications. However, there are no such design guidelines available in the literature to generate a radial rotor blade especially for supersonic flows. The conventional 1D preliminary design methods, generally available for radial rotor blades, are not very reliable in supersonic flow regime. This calls for a novel design philosophy for supersonic radial rotors that accounts for supersonic flow phenomenon early in the design phase to ensure efficient turbine performance. The objective of this thesis is to put forth a design methodology for supersonic radial rotors that acknowledges the flow equations in the design procedure. The proposed design ideology is to extend the supersonic axial rotor design procedure by including the fictitious forces' effects in the governing equations to generate a radial rotor using MOC. These pseudo forces (Centrifugal & Coriolis) are experienced by the flow in the radial runner if seen from the rotor's rotating frame of reference. Furthermore, the dense gas effects are also included in the design using the properties from CoolProp library. A design tool is developed in PYTHON that implements the aforementioned ideology in the design procedure. This thesis takes the first step towards developing a fully functional design methodology for supersonic radial rotors. Therefore, a simple case of a constant force's influence on the rotor geometry is studied in this work and compared with the CFD simulations in both perfect & dense gas region. Future research will focus on understanding the complex behaviour of organic fluids in the presence of an external force and modify the design procedure accordingly.

Organic Rankine Cycle systems are gaining increasing attention as a modular, cost-effective and decentralised thermal energy recovery solution. The design of an efficient expander in the ORC system is crucial to its rapid uptake by the industry. Compared to volumetric expanders, a radial turbine provides advantages in terms of flexibility and better part-load performances. The design of ORC radial turbines is faced with two challenges. Firstly, the expansion in the stator occurs close to the vapour-liquid dome and the critical point, where the ideal gas assumption is no longer valid. Such expansions fall under the Non-ideal thermodynamic regime. Secondly, due to the lower speed of sound in the working fluid, the expansions are supersonic and are accompanied by compressible flow features such as shock waves, expansion fans which generate entropy and reduce the performance of the expander. There exists significant validated design methodologies and empirical relations for conventional turbomachinery. However this is not the case for unconventional turbomachinery where established loss correlations are not applicable. Hence, there is a knowledge gap in validated design tools for unconventional turbomachinery like ORC radial turbines. The thesis forms a part of the validation campaign of the open-source flow solver, SU2 and approaches the exercise by studying expansions in two paradigmatic test cases, namely (i) a linear stator cascade and (ii) a converging - diverging nozzle. The study of expansions through a linear cascade is a preceding step to the study of a rotating radial turbine. Consequently, a RANS simulation on the single channel blade passage is studied under the validated assumption of flow periodicity. The expansion takes place from inlet total conditions of 18.4 bara and 525 K (Z =0.558) to a static pressure of 1.95 bara (Z = 0.951) and exit flow of Mach 2. Two types of response quantities are studied namely direct and system response quantities. Direct response quantities are those that can be measured directly such as the pressure, Mach, density. System response quantities are derived from direct measurements and provide information to characterise the performance of the stator. The selected system response quantities are: (i) Pressure loss coefficient (Cp= 0.074), (ii) Kinetic energy loss coefficient (ζKE= 0.115), (iii) Entropy loss coefficient (ζs=0.118), (iv) Base pressure loss coefficient (Cpb= -0.065), (v) Standard deviation of exit flow angle (σβ2 = 1.244) and (vi) Standard deviation of exit Mach (σM2= 0.033). Experimentally, it is possible to measure the pressure loss coefficient, base pressure loss coefficient and the flow uniformity at the outlet using a combination of pressure and direct velocity measurements. Through a Design of Experiments approach, the sensitivity of the flow to input uncertainties was studied through a stochastic collocation based forward propagation of the uncertainty. The input uncertainties considered are the inlet total pressure, fluid viscosity and critical point properties. The Sobol indices from the uncertainty quantification indicate the more dominant influence of the critical point properties over other inputs considered. The results also validate the use of a constant viscosity assumption for the RANS simulation. The subsequent planned unit test case was to characterise the expansion through an optimised stator blade row. To this end, a deterministic adjoint based optimisation was performed with the objective function of minimising entropy generation. The resulting geometry was studied and resulted in a pressure loss coefficient improvement of around 4%, although stator flow uniformity at the exit was compromised. The uncertainty quantification performed on the optimised blade geometry yielded robustness improvements on the pressure loss coefficient and reduction in uncertainties associated with direct response quantities, although no strong correlations can be drawn between the improvements in uncertainty and the deterministic optimisation. Given current machining tolerances, the realisation of such a negligible geometry change is not viable from an experimental perspective. The second section of the thesis deals with experimental investigation of expansions in a converging-diverging nozzle through a matrix of isobars with increasing degree on non-ideality. Two isentropes with inlet pressure of 2.73 bara (Z = 0.9526) and 6 bara (Z = 0.901) with pressure ratio of 8.76 were performed with recording of pressure, temperature, flowrate and density measurements along the ORCHID. The flow field was visualised using the schlieren method using a z-type layout. The thesis reports the first measurements of total pressure before the nozzle inlet and vapour density and flowrate measurements. The experimental data was post-processed to identify steady-state and quantify the Type A and Type B uncertainties. The schlieren images were used to extract information on the Mach distribution along the nozzle mid-plane using an inhouse line extraction tool. Lastly, a 2.5o wedge at the exit of the nozzle is used to generate oblique shock waves that are then manually measured. The flow field at the exit of the nozzle is in the ideal region where the shock angle only depends on upstream Mach number. Hence, the experimentally observed oblique shock angles for the off-design case are close to on-design (Inlet pressure = 18.4 bara) predictions.

This document contains the final Master Thesis Report to obtain the Master of Science on Aerospace Engineering (Flight Performance & Propulsion - Propulsion & Power) at the Delft University of Technology. The research work focuses on characterizing the interaction of a radial inlet turbine with a downstream diffuser through the size of the rotor tip gaps. In the implementation of power turbines it is common to add a diffuser downstream of it to lower the rotor exhaust pressure and thus increase power extraction for the same inlet conditions. However, diffusersare bulky and take a lot of space in the assembly. This lowers the power density of the machine, increases installation weight in transport applications, and installation costs in ground based operations. In the quest to obtain compact diffusers, researchers noticed that the non-dimensional static pressure recovery (Cp) of this device was higher when operating downstream of a turbine than with uniform inlet conditions. The publications in this field are relatively scarce, and thus there is a knowledge gap with immediate practical application. Some researchers have focused on the interaction of turbine vortical structures with the boundary layer of the diffuser. They have found out that under this conditions the boundary layer can support steeper pressure gradients without detaching. This is only applicable in very steep diffusers that would stall in isolation. Notably, all the publications in this field deal with axial machines, and as the text will show the picture changes considerably when applying the theory to radial turbines. This work studies another side of the problem. The text will focus on stable diffusers, and thus there is no boundary layer that needs reinforcement. Turbine rotor tip gaps generate an increase of entropy and reduce turbine power generation. However, these gaps also cast powerful vortexes that affect the diffuser flowfield. This project will study the effect of the tip gap configuration on the pressure recovery of the diffuser, in order to better understand the trade-offs and support the design process. Tip gap sizes are usually determined from mechanical constraints. This work provides insight into the real cost of a tip gap by analyzing the assembly turbine + diffuser, and thus it guides the designer into where to focus his efforts when working with these mechanical constraints. The general conclusion of the project is that there is a range of tip gap sizes where the performance of the diffuser in enhanced. In this range, the performance cost in the rotor for increasing the tip gaps is partially compensated by a better pressure recovery in the diffuser. Furthermore, it has been discovered that in a radial turbine it is more influential to close the leading edge gap (axial gap), even when this implies opening the trailing edge (radial) one. This information will guide engineers when choosing the bearings for the machine and dealing with design trade-offs. As a byproduct of this research, a novel study in radial turbine flow structures is carried out. Not a lot is known about the vortical structures in these machines, and this work provides a first qualitative description about origin and behaviour of this vortexes. Furthermore, different flow modes are identified and correlated to a simple geometrical parameter that is readily available in 0D designs. The combination of this information with previous tip gap flow models, and diffuser-vortex interaction models, will allow the obtention of improved losses models including the effect of the diffuser from the conceptual design stage. This data could also feed into noise models for radial turbines, a novel field of interest. This work not only provides useful design guidelines, but also sets down the basis for future research leading to a better understanding and modeling of radial inflow turbines.

With an increasing complexity of systems, systematic design space exploration is required for fair analysis and comparison of architecture design options. However, there is still a need for a benchmark problem to support the development & evaluation of optimization algorithms and system architecture design space modeling methods, and to educate stakeholders in system architecture optimization. The thesis focus lies on the creation of such benchmark problem with an application to aircraft jet engine design, which is subsequently tackled by developing an MDO tool using a system architecting approach with mixed-discrete and multi-objective capabilities. The tool was built with the open-source software pyCycle and OpenMDAO and can generate, analyze and compare different aircraft engine architectures on several disciplines including engine thermodynamic cycle analysis, weight, geometry, emissions and noise. Therefore, it is able to tackle the black-box, mixed-discrete, multi-objective and hierarchical nature of system architecture optimization problems.

Curved exhaust ducts are used in aero engine applications for different purposes, including thrust vectoring, shielding of parts from the exhaust or improving stealth properties. Their integration, however, regularly represents a design problem due to flow separation and high aerodynamic losses occurring in the bend. Curved duct flows for both compressible and incompressible conditions have been studied extensively in the past. However, no experimental results of a high Reynolds number flow through a turbine exit guide vane (EGV) followed by a curved duct have yet been published. An in-depth CFD analysis of the aerodynamic effects is therefore carried out, using the RANS solvers TRACE (at MTU Aero Engines) and SU2, to analyze the flow field and geometric sensitivities of a high Reynolds number flow through an EGV followed by a 90 degree bent duct. The geometry of interest is investigated at a Reynolds number of Re=10^6 and a ratio of bend radius to duct diameter of one. The inflow conditions are prescribed to closely resemble typical flow conditions at the low pressure turbine exit plane of a turbojet engine. After validating the solver with experimental data using the test case of a 90 degree bent duct, an initial CFD analysis of the combined exit guide vane geometry with curved duct is carried out to identify the dominant flow phenomena and mutual effects of the EGV and the bend. Three zones of flow separations are found, each at the convex and concave sides of the bend and one at the lower side of the plug. Secondary flows caused by the bend are found to have an effect on the flow upstream of the EGV, leading to a non-uniform flow inlet angle and aerodynamic loading of the individual blades. Separation downstream of the EGV is influenced by the presence of low velocity wakes from the EGV. Subsequently, a sensitivity study is carried out to find the effect of different geometrical parameters on the flow field. Main investigated parameters are the ratio of bend radius to diameter (R_c/D) and the distance between EGV and bend (l/D). Additionally, the aerodynamic effects of the circumferential EGV positioning, swirl and the plug shape are investigated. It is found that an increase in both bend radius and distance between bend and EGV improves aerodynamic efficiency, while swirl can decrease pressure losses in the duct for small bend radii of R_c/D<0.8. Improving the plug shape and rotating the EGV allows to further increase the aerodynamic efficiency without weight increase. For further optimization of the geometry, it is recommended to include duct geometries with a non-constant bend radius and outer duct diameter to increase the design space.

A generalized isentropic gas model is derived following earlier work by Kouremenos et al. in the 1980s, by replacing the traditional adiabatic exponent γ by the real exponents γPv, γTv, and γPT, describing the isentropic pressure-volume, temperature-volume, and pressure-temperature relations respectively. The real adiabatic exponents are expressed as functions of state variables to take into account compressibility effects on the isentropic behavior of substances. Due to the implicit analytical nature of the real exponents, any equation of state or thermodynamic library can be used for their evaluation. The theoretical limits and overall behavior of the real isentropic gas model are explored for a Van der Waals substance. In the two opposing physical limits, the model is shown to reduce to the incompressible substance model for liquid densities and the ideal gas model as the temperature increases or the pressure goes to zero. The relation of the generalized isentropic gas model with other thermodynamic properties is explored, leading to the development of specific heat relations and other thermodynamic properties in terms of the real exponents γPv, γTv, and γPT. Besides providing alternative schemes for their evaluation, special features of thermodynamic properties such as the state of maximum density and inversion temperature may be related to the value of the isentropic exponents determined by the local compressibility of the substance. Due to the exact definitions of the real adiabatic exponents at a state point, the relations between properties is thermodynamically consistent – another physical requirement. The generalized isentropic gas model is then applied to isentropic flows to derive traditional gas dynamic relations such as speed of sound, stagnation properties, and choked flow conditions for non-ideal compressible fluid flows. Exact solutions are provided for Prandtl-Meyer expansion fans, and approximate Rankine-Hugoniot jump conditions are explored for real gases. Finally, attributes of the fundamental derivative of gas dynamics are explored under the generalized isentropic gas model to gain new insights into its mathematical properties. Under the generalized isentropic model, the fundamental derivative is shown to satisfy both liquid and gaseous physical limits. Non-classical behavior is attributed to higher-order derivatives of the real exponents. The application of the generalized isentropic gas model is demonstrated and validated for use in non-ideal compressible fluid dynamic (NICFD) codes by simulation of the one-dimensional Euler equations for a standard shock tube problem. Several numerical schemes for the evaluation of thermodynamic properties of varying levels of accuracy are presented for evaluation of the isentropic gas model. In the application of the shock tube problem, the general equation for the speed of sound is demonstrated to be equivalent to the speed of sound of a Van der Waals gas, proving the validity of the model.

Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. Modeling the impact of secondary path and leakage flows on the performance and axial thrust of the pump is critical during the early design phase. Previous studies derived simplified models and empirical correlations to model leakage loss and determine the pressure distribution in the side gaps. The prediction capabilities of these models are subject to limitations and uncertainties mainly due to the interaction between impeller exit flow and sidewall gaps especially at partload operation. The effects of the leakage injection on the impeller flow are not fully understood and not captured by analytical correlations from literature. These can have destabilizing and detrimental effects on the hydraulic performance and the head curve, leading to a divergence betweenmodel predictions and experimental results. In this project a reduced order model is developed to capture the effects of the sidewall gaps on turbopump performance and determine the axial thrust on the rotor. High fidelity calculations of the pump and volute are performed to identify the key physical mechanisms associated with loss generation in leakage paths and improve upon existing analytical models found in literature. Axial thrust and leakage losses are modeled numerically with a single passage geometry to capture the interaction of impeller flow and leakage flow through the side gaps at multiple operating points. It is found that the recirculation of pressurized fluid, leakage mass flow, is the most impactful effect on the performance. The volumetric efficiency is captured within 16% with the analytical model and 6% with the single passage numerical calculations at nominal operation. At partload operation however, the accuracy of both models reduces. Disk friction is under predicted by the investigated analytical model and axial thrust is over predicted. For both effects the reduced numerical model shows promising results that improve the prediction capability around nominal operation and offer a higher flexibility required for the early design phase as compared with the high fidelity, full annulus calculations. For the investigated cases the leakage injection to the impeller flow does not show a destabilizing impact on the characteristic behavior.

There is an ever-growing need for clean renewable energy in modern day society. Currently, major developments are being made on Decentralised Power conversion systems that use renewable energy sources. The application of ORC radial-flow turbines in Decentralised Power conversion applications seems promising, however, application of this technology is still sparse due to the lack of knowledge of the design, resulting in an extensive duration of the design cycle. The purpose of this thesis, is to reduce the duration of the design cycle, which can lead to a faster transition towards Decentralised Power Generation to help meet the increasing energy demands.

Direct design optimisation methods for turbomachinery blades have consistently gained in popularity over time. This has been accomplished by significant advances in optimisation algorithms, and notably the development of adjoint-based gradient computation, which provides objective sensitivity information at comparatively low computational cost. This enables the use of first-order optimisation methods for more complex design problems. On the other hand, interest in inverse design methods has steadily diminished over time. This is because these methods come with significant disadvantages or weaknesses, such as requiring strong simplification or only being narrowly applicable, e.g. incompressible, potential flow. However, by applying the advances in optimisation techniques to an inverse design strategy, a versatile and effective design tool can be created with functionalities not offered by direct design techniques. In order to evaluate the feasibility and usefulness of such a method, an adjoint-based inverse design tool for 2D axial compressor blades was developed. The tool requires a baseline geometry and a target pressure or isentropic Mach distribution along the blade wall. It then analyses the flow field around the initial geometry and modifies autonomously the shape to reach closer to the desired target distribution. A discrete adjoint approach is used to compute the gradients of the objective with respect to each of the design variables. The objective and its gradients are used by a the first-order SQP optimisation algorithm to iterate the design until it matches the target distribution as closely as possible. The newly developed design tool was used in an attempt of improving the performance of the NASA stage 35 stator blade. For comparison, also single- and multi-objective direct design optimisations were performed. It became clear that while significantly cheaper in computational cost and simpler in program complexity, the inverse design method cannot be used to reliably produce improved designs. The value of the resulting design is entirely dependent on the quality of the target distribution. It is up to the designer to create a target which is physically possible and leads to increased performance. Still, inverse design does provide unique capabilities not offered by direct design. The tool was successfully used to retrieve the blade geometry used in a different work based on published isentropic Mach number distributions. The obtained geometry closely matched the actual geometry used in the source work. This demonstrates the accuracy of the method and the fulfilment of a use case not offered by other design methods.

Exploring and utilising space to benefit humankind is undeniably a costly venture where the higher the mass of the satellite, the higher the cost to launch. However, the rewards of space-related achievements are indisputable. These achievements range from increasing our understanding of the universe using satellites, such as the Hubble Space Telescope,to improving the predictive ability of weather, climate, and natural disaster phenomena with the help of Earth observation satellites.@en

In 2022 the aviation sector accounted for 1.9%of global greenhouse gas emissions, 2.5% of global CO2 emissions, and 3.5% of effective radiative forcing. To reach the long-term target of net zero emissions, revolutionary aircraft designs, featuring electrified or hydrogen powered propulsion systems, are needed. At the same time, the electrification of the non propulsive aircraft subsystems is necessary to comply with the requirements of emissions abatement in the short and medium time horizon. Among the auxiliary subsystems, the Environmental Control System (ECS) is the largest consumer of non-propulsive power, accounting for up to 3-5% of the total fuel burn. The replacement of the conventional Air Cycle Machine (ACM) with an electrically-powered ECS based on the Vapor Compression Cycle (VCC) system could enable: i) a substantial decrease in fuel consumption; ii) a finer regulation of the relative humidity in the air distribution system, leading to improved air quality in the cabin and flight deck; iii) a reduction in maintenance costs and an increase in system reliability, due to the removal of the maintenance-intensive bleed system. However, the adoption of VCC systems in the aerospace sector has been historically very limited, due to safety concerns regarding the ozone depleting potential, toxicity and flammability of the working fluids used as refrigerants, as well as because of a lack of research specifically targeting airborne applications. This dissertation documents research work performed as part of the NEDEFA project, which entails the investigation of VCC-based ECS architectures powered by oil-free highspeed centrifugal compressors. The first objective is to advance of the state-of-the-art regarding high-speed compressors operating with gas bearings, i.e., the key technological enablers of airborne VCC systems. The second target is to develop of a methodology for the integrated design of aircraft ECS, namely, a design philosophy in which the system and the main components are optimized simultaneously. The main outcomes of this work are the development of a preliminary design model for high-speed compressors, extensively validated with experimental data and computational fluid dynamics simulations, and the implementation of an integrated design framework for aircraft ECS, embedding a multi-point and multi-objective optimization strategy. The compressor model has been applied to derive design guidelines for single-stage and twin-stage machines operating with arbitrary working fluids, as well as to perform the fluid dynamic design optimization of the compressor that will be installed in the IRIS (Inverse organic Rankine Integrated System) test rig of the Propulsion and Power Laboratory. Furthermore, the integrated design method has been used to size and compare the performance of two alternative ECS configurations for a single-aisle, short-haul aircraft resembling the configuration of an Airbus A320, i.e., a bleedless ACM and an electrically driven VCC. The results reveal that the optimal VCC system could be both more efficient and lighter than the corresponding ACM architecture, leading to potential fuel savings in the order of 20% for the prescribed application. @en

In this thesis the goal is to grow the fundamental understanding of the Shrink Sleeve Labeling process, speciffically on the steam tunnel equipment. The research is focused on the heat transfer phenomena of the ow on to the bottle. From literature was found that the problem could be split into 3 general domains: 1) free jet ow, 2) heat transfer by impinging jet and 3) condensation heat transfer. The condensation part was seen as an optional part since it would have a big impact on the way the numerical model will be setup. This thesis therefore is setup to handle the free jet ow and heat transfer by an impinging jet.

Climate change is a pressing issue. Hydrogen aircraft are an excellent alternative to current kerosene aircraft, and appear to be the most viable long-term solution for net-zero flight. One of the main challenges for hydrogen aircraft is to sustain a low boil-off rate. This research explores the solution of active cooling of the LH2 mixture in the fuel tank. It focusses on the construction of a thermodynamic model for cryogenic liquid hydrogen fuel storage in aircraft, a system model for single-stage Reverse Turbo-Brayton Cryocoolers (RTBC), and the conceptual design of the RTBC’s miniature high-speed compressor. This is used for the integrated modelling of the RTBC, compressor and LH2 fuel tank for an exploration study of the carbon neutral hydrogen concept of the long-range Flying-V aircraft. Results show that the RTBC might offer a valuable addition to the Flying-V design space for boil-off control in addition to careful design of the insulation and tank shape.

The progressive replacement of traditional, fossil fuel-based energy transformation processes through a worldwide increase in the usage of alternative, renewable primary energy sources has boosted interest for thermodynamic engines operating with fluids other than steam, air and flue gases. Especially, the availability of low-quality (i.e. low temperature) heat sources, as it is the case concentrated solar systems, geothermal and biomass plants, makes use of classic working fluids, such as water, no longer viable. Similarly, efforts to move to delocalized, small-capacity (about 0.1–10 MW) power plants and mobile powertrains still running on C-based fuels but with a virtually vanishing C-footprint as backup for volatile renewable sources have brought renewed attention to waste heat recovery applications based on unconventional technologies involving non-ideal fluid flows mostly centred on the concept of the organic Rankine cycle (ORC) and of supercritical CO2 (sCO2) technology. Such systems operate with working fluids (e.g. siloxanes, refrigerants, CO2) and in thermodynamic states exhibiting thermo-physical behaviour largely departing from that of an ideal gas, often close to the critical point, either in vapour or in two-phase conditions. Though technologies exploiting the properties of non-ideal flows find already a widespread application in space propulsion (rocket engines) and in the oil and gas industry, their relevance and popularity are also growing in residential applications especially due to the advent of the next generation of heat pumps. Whilst representing an attractive evolution of energy transformation processes, the migration to trans- or supercritical and generally unsteady operations poses new challenges for the design, optimization and maintenance of systems and their components. The predictability of the performance and life cycle of individual components requires in particular an accurate description of the thermodynamic and transport properties of the working fluid, on one side, and a faithful modelling of all relevant flow features on the other side, especially in combination with turbulence, compressibility and possibly phase change effects, as it is the case in turbomachinery (compressors and expanders). @en