Pulse Detonation Engines (PDEs) are propulsion systems that use repeated detonations to generate thrust. Currently in early stages of development, PDEs have been theorised to have advantages over current deflagration based engines. Air-breathing PDEs could attain higher specific impulse values and operate at higher Mach numbers than today's air-breathing engines, while Pulse Detonation Rocket Engines (PDREs) could become a lighter, cheaper, and more reliable alternative to traditional rocket engines. There are still however, many technological hurdles to overcome before PDEs can be developed into practical propulsion systems, one major barrier being management of their immense heat loads. This thesis outlines the development of a numerical model for determining thrust performance and heat load characteristics of PDEs. The model is based on a set of analytical equations which characterise the gas dynamics inside the engine throughout it's cyclic process. Being numerically light -when compared to CFD analysis- the model allows for fast turnaround of results and the ability to sweep through parameters to determine optimum operating conditions to maximise engine performance and reduce heat load. In this study, the working principles of the model are described and it's outputs are validated against data from published experimental and numerical studies. The model is then used to conduct a comprehensive parametric study on the effects of various reactant combinations, operating conditions, and engine geometries on engine thrust, specific impulse and heat load. Lastly, a brief study is conducted on the feasibility of regenerative cooling for PDEs, using model outputs to determine if a heat balance can be achieved and the performance losses and complications that would result.

This thesis considers vibration drilling in aerospace materials. The tolerances of the drilled holes in aerospace industry are very low since hole quality is an important factor. Conventional methods of drilling create long ribbon formed chips that increases the heat formation and decreases the hole quality. The solution is to introduce low frequency axial vibrations at the drill to break the chips. Smaller chips are easier to evacuate and leads to less heat formation and increased hole quality. Mechanical solutions to create the axial vibrations are commonly used in so called Advanced Drilling Units (ADUs). These drilling machines mounts on the surface to be drilled, actuates the drill with a feeding motion and drills the hole automatically. The ADU PFD1100 from Atlas Copco uses a mechanical chip breaking module called the ChipLet. The ChipLet has fixed amplitude and frequency hence the ChipLet module needs to be replaced to change vibration parameters. This thesis proposes the idea to instead use an electromechanical solution to create the vibration with an electric motor. This would make the possibility to change vibration parameters on the fly. A motor has been coupled to the feeding mechanism of the PFD1100 using a created prototype. The motor modifies the feeding motion of the spindle by doing a superposition of the constant feeding with a sinusoidal motion hence creating vibrations on the spindle. To compare the prototype to the current ChipLet, force and torque data have been gathered and analysed with spectral signal processing using the power spectral density estimate (PSDE). Conventional drilling with both the prototype and the ChipLet have been tested and analysed to use as a baseline and reference. The prototype shows that electronic control of the vibration parameters is possible. The prototype is also capable of breaking the chips although at lower frequency and amplitude than the ChipLet. The PSDE shows that conventional drilling reference frequencies are present in the vibration drilling data sets. The PSDE also shows that both the prototype and the ChipLet have several overtones in addition to the main harmonic.

The purpose of this master thesis is to provide a reliable methodology to predict the forced response of a monoblock bladed drum from a low pressure compressor. Pre-test forced response calculations have already been made at Techspace Aero in 2013. Now that experimental data are available, the methodology has to be adapted to ensure the best numerical-experimental correlation possible. The final goal is that, at the end of the thesis, engineers at Techspace Aero will be able to launch reliable forced response simulations within a short amount of time. For the sake of confidentiality, some data are not revealed, such as the engine name, some numerical values (forced response, aerodynamic damping, frequency of the mode etc…) and axis scales. In this paper, the study focuses on the forced response of a rotor blade from the first stage under the excitation from the upstream stator. The mode under investigation is the 2S2, the one that responded during the experiment. The TWIN approach is used to compute the forced response of the rotor blade. With this approach, a steady stage computation has first to be carried on as an initialization. Then two unsteady computations are necessary. The first, without blade motion, will provide the excitation aerodynamic forces. The aerodynamic damping will be extracted from the second one, where the motion of the blade is imposed on a given eigenmode. The forced response can then be computed with these two results and some additional structural data.  The results will be compared to the experimental value.

A parametrical study is performed of coolant dump gas ejectors for a sandwich rocket nozzle design. Five geometrical variations are simulated in four ambient conditions (static, subsonic, supersonic, vacuum) using an in-house CFD solver. The test cases are compared with a baseline case and the resulting thrust and ISP are evaluated on a local and global level. A longer dump wall is found to give the best performance in all ambient cases, with a second possibility of reducing the circumference of the nozzle end stiffener. The possibility of post-ejection coolant gas combustion is encountered for high ambient pressure, high subsonic velocity flight.

Modelling condensation of water vapour is important in a number of engineering applications, such as nuclear reactor containment, rocket engine nozzles and heat exchangers. The current study investigates the possibilities of modelling condensation induced by a cold surface in a flow at high velocity and temperature. A number of non-condensable gases are present in the flow. The possibilities of condensation modelling are investigated in ANSYS CFX and ANSYS Fluent, with focus on ANSYS CFX. A case study is done of a 2D flat plate, with water vapour and non-condensable gases at varying temperatures and velocities. The condensation models in ANSYS CFX are investigated for a few basic flow setups and the model deemed most appropriate for wall condensation is investigated in greater detail. The wall condensation model in CFX is investigated to a greater extend, and compares well with an analytical solution for laminar flow. The complexity of the flow is gradually increased to determine limitations and best practise settings for flows at high velocity and temperature. For isothermal walls and for a conjugated analysis, using a solid with a specified adiabatic wall temperature and heat flux coefficient to induce condensation, the wall condensation model works well for grids with a y+ above 1. For finer grids, convergence is found to be difficult to achieve. The choice of material properties for water vapour was found to play an important role in terms of stability. Real gas properties to define the water vapour material properties are deemed important to avoid unphysical results in terms of the temperature. The wall condensation model in ANSYS CFX is deemed to be an appropriate choice for future work with respect to validity and reduced complexity. If the wall condensation model in ANSYS CFX does not prove adequate, it is recommended to investigate an Euler-Euler multiphase model in ANSYS Fluent with condensation and the Eulerian wall film model enabled.

On the one hand, airlines order new planes and the worldwide fleet increases, while, on the other hand, the market pressure, the rise of fuel prices and other factors contribute to regular changes in the technology. These drivers may impact maintenance activities and support to operators, and the number of issues occurring on in-service aircraft. In-service and production queries are a specific type of support activities followed-up by propulsion systems integration engineers from the aircraft manufacturer, such as Airbus. These technical questions can address any of the engine’s systems and must usually be answered to within a short timeframe as they might delay a flight or the delivery of an airplane. In the global scope of knowledge management inside the company, these engineers realized their loss of not capitalizing these activities and promoted this project. An adapted application has been developed to share the experience among programs and support the engineers for the treatment of such queries. As the focus of the project was put on assessing the actual need of the future users to provide an adapted tool, the database should prove its performance over the long term. This paper details the different steps of the project: analysis of the need, specifications, programming and testing, that led to meeting this specific need for capitalization. / <p>Presentation via videoconference for distance-based students</p>

This thesis presents a tool for conceptual design of a traditional configuration aircraft by using a parametric six degrees of freedom (6DOF) flight mechanics model implemented in the Modelica Language using Wolfram System Modeler. Being first only able to model and simulate the uncontrolled flight of an aircraft with fixed mass and centre of gravity (CG), and requiring detailed aerodynamic parameters as an input, the 6DOF model is improved by developing new features to reduce the number of required inputs while also increasing the data output of the simulations. First, the propulsion submodel is added with models for alternative propulsions to the existing model of turbofan engines. The energy and fuel consumption is also modelled for all propulsion types, and thus the aircraft model has no longer fixed mass properties, except for aircraft with electric propulsion. To further evaluate the fuel consumption for pre-defined flight missions with given flight speed, altitude and track angles, autopilots for a few different aircraft types are developed. Additionally, the 6DOF model is improved by establishing algebraic and statistical relationships between the aircraft geometric input parameters, aerodynamic coefficients and moments of inertia such that the values for the two last mentioned can be estimated inside the 6DOF model based on the minimum amount of design variables, geometric input parameters and aerodynamic properties of the 2D airfoils used in the wings. Ultimately, the improved 6DOF model is evaluated and analysed in terms of its performance in initial weight estimation on aircraft conceptual design stage as well as in predicting the aerodynamic properties.

Hybrid rocket motors have become of great international interest during the last couple of years. A hybrid rocket motor is propelled by the use of a solid fuel and a liquid oxidizer. The fundamental principle of the hybrid propelled system is that the liquid oxidizer is injected into a combustion chamber to enable the combustion of the solid fuel. The exhaust gases are then accelerated through a nozzle to supersonic velocity to produce the desired level of thrust. To describe the overall performance of a propulsion system, it is common use the specifc impulse which expresses the performance as the total impulse per mass unit of propellant. However, in order to optimize a propulsion system, it is necessary to consider the entire system with the oxidizer tank, feed system, combustion chamber and nozzle. The issue with using the specifc impulse as a performance index is that it does not consider the total mass of the propulsion system. Therefore, this thesis will instead analyze the system specifc impulse, which expresses the performance as the total impulse per mass unit of propulsion system. By studying the entire hybrid propulsion system it is possible to determine the relations between the various parameters of the diferent components and should therefore be able to optimize the mass, volume and system specifc impulse of the system. This master’s thesis aims to illustrate how the hybrid propulsion system can be optimized depending on various fxed parameters. This analysis studies a generic hybrid propulsion systemwith Hydroxyl-terminated polybutadiene (HTPB) as a solid fuel with diferent combinations ofoxidizers. Each oxidizer- and fuel confguration shall have identical combustion chamber presssures and shall generate the same total impulse. Nevertheless, each combination will result indiferent specifc impulses since the optimal confguration for each combination will generate diffferent oxidizer and fuel masses. It is then desirable to analyze how the diferent components ofthe propulsion system are affected by the required oxidizer and fuel for each optimal confgurationand how it drives the design of the system and generates diferent system specifc impulses.

In this degree project in aerospace engineering, a preliminary design of a UAV (Unmanned Aerial Vehicle) was performed. The UAV was intended to be used as a complement to the Swedish maritime administration’s helicopters, which cannot operate under limited visibility conditions. Its main mission would consist of surveillance. The UAV was therefore designed for some series criteria that were based on the customers’ requirements. The primary literature that was used was John D. Andersons Aircraft performance and design. Otherwise, historical statistical data from other aircraft was used and numerous assumptions were made. The result was a relatively small UAV named The Condor, weighing 25.6 kg with a wingspan of 2.5 m and operational in an altitude of 3500 m with a cruise speed of 81 knots. The UAV’s range is 70 nautical miles and is also able to operate in up to six hours. It should be able to manage a 300 m long runway. The chosen wing profile was the NACA 1412 with a maximal thickness and camber of 12 % and 1 % of the chord length, respectively. As for the stabilizer, the symmetric wing profile NACA 0012 was chosen. A so called constraint analysis was performed in order to determine the engine choice and thewing loading. The chosen engine was a 3.1 horsepower piston engine provided by Ricardo. The dimensions of the fuselage were designed only to fit the payload and no detailed analysis was done. It became 2.3 m long and with a maximal diameter of 0.3 m. / I det här kandidatexamensarbetet i flygteknik gjordes en preliminär design av en drönare. Drönaren skulle användas som komplettering till sjöfartsverkets helikoptrar, som inte kan användas vid mycket begränsad sikt. Dess huvudsakliga uppdrag skulle bestå av övervakning. Drönarens utformades därför efter en rad kriterier som baserades på uppdragsgivarens krav. Den huvudsakliga litteraturen som användes var John D. Andersons Aircraft performanceand design. I övrigt användes historisk statistisk data från andra flygplan och ett flertal antaganden gjordes. Resultatet blev en relativt liten drönare som döptes till The Condor och fick en vikt på 25.6 kg, med ett vingspann på 2.5 m och som opererar på 3500 m flyghöjd med en marschfart på 81 knop. Drönarens räckvidd är 70 nautiska mil och den kan därutöver operera i upp till sex timmar. Den bör klara av en landningsbana på 300 m. Som vingprofil valdes NACA 1412 med en maximal tjocklek och camber på 12 % respektive 1 % av kordalängden. För stabilisatorn valdes den symmetriska profilen NACA 0012. En så kallad ”constraint analysis” genomfördes för fastställande av motorval och vingbelastning. Motorn som valdes blev en 3.1 hästkrafters pistongmotor från Ricardo. Flygplanskroppens dimensioner utformades endast för att få plats med nyttolasten och ingen noggrannare analys genomfördes. Den blev 2.3 m lång med en maximal diameter på0.3 m.

The objective of this master thesis was to investigate an ablative liner for the T-Minus DART booster that will accelerate a dart to Mach 5.2 within five seconds. An oxyacetylene torch test was used to sort out the obviously bad materials. Glass fiber/epoxy, with and without alumina as fire retardant, and carbon fiber/epoxy were selected for further investigation. A sub-scale motor was built to expose the materials for conditions similar to the booster conditions in regard to temperature, chemistry, flow velocity and pressure. The target pressure could not be reached in the sub-scale motor but a polynomial function was fitted to the data in order to extrapolate the data and estimate the ablation rate at 7 MPa. The final design is always based on measurements on full scale motors. This could not be done within this report. Recommendation for future work is to use an insulator of 1.8 mm of carbon fiber/epoxy or 1.3 mm of glass fiber/epoxy/alumina for the sub-scale firings to come.