In this work an overview is provided of various rocket thrust frames found in open literature and of various existing single relation low-level models for rocket thrust frame mass estimation. Based on data obtained from literature, a new relation has been set up that supports making a distinction between three main types of thrust frames identified in this work. Some possible improvements have been pointed out that can lead to better low-level models. In this work also foundational work has been performed towards an intermediate level model for thrust frame estimation. The resulting model provides clarity on how various important parameters like thrust, engine mass, acceleration level, frame dimensions, materials, etc. affect thrust frame mass. The accuracy of the results obtained could not yet be proven to agree with the required range for intermediate level models, but points for improvement have been listed

In this work two Vehicle Equipment Bay (VEB) mass estimation models are applied to the estimation of the mass of the Ariane 5G VEB. The two models presented include a low-level (level 0) and an intermediate level (level 1) fidelity model. The low-level model provides a Most Likely Estimate (MLE) of 1118 kg compared to a "true" value in range 1430 to 1500 kg. The intermediate level model provides an MLE of 1506 kg and shows a clear improvement in the estimated mass as compared to the true mass. It is advised though to further investigate the validity of the two models used with focus on the intermediate level model as it seems better capable of providing the required accuracy.

This project (poster) explores and maps transdisciplinary skills in the TU Delft curricula and challenge based education. Courses to address these skills have been identified by means of keyword search in the course descriptions. Interviews are used to explore the transdisciplinary approaches addressing reasons, values, learning activities, assessment and professionalisation. The exploration was initiated by a multidisciplinary group of educators from different TU Delft faculties. The initiators noticed that transdisciplinary skills are regularly part of a hidden curriculum, delicate to define or grasp, bear different names, are rarely made explicit or maybe even are considered a taboo. As such, the transdisciplinary skills remain unspeakable.

The exploration is made within the Technical University of Delft. It is to be expected that lessons learned will not be exclusive to this context and can be applied in other settings that aim for societal impact of science and education as well.

Higher educational institutions have broadly adopted Collaborative Engineering Design (CED) activities to prepare students for complex problem-solving in multidisciplinary settings. These activities are non-linear and mediated by various social practices and tools. Therefore educators might struggle in facilitating the achievement of specific learning goals. Embodied cognition is an approach that explains non-linear behaviour through orgamism-environment interactions and might therefore provide educators with insights on how to prompt students towards desired actions in CED activities. According to embodied cognition, we learn through actions that emerge as a response to a problem (task) and environmental constraints. Educators can guide students’ behaviour by proposing tasks and adapting the environmental constraints of a learning situation, thus creating a field of promoted action. In this paper, we outline the progress of a design-based research in which insights from embodied cognition are implemented to promote desired student behaviour in CED activities. We report on the results of our problem-exploration phase. A systematic literature review and focus groups with students revealed that students are often hesitant to adopt new practices and tools that could potentially improve their collaborative design process. Next, we propose three theory-based design principles in which the task and environmental constraints are leveraged to foster the adoption of practices and tools and apply them to CED activities. Finally, we will share preliminary observations of the learning processes triggered by the designed activities and outline the directions for future research.

Contribution: This article presents a comprehensive overview of characteristics of educational designs of collaborative engineering design activities found in literature and how these characteristics mediate students' collaboration. Background: Engineers have to solve complex problems that require collaboration. In education, various collaborative engineering design activities have been implemented to prepare students for these professional practices. According to cultural historical activity theory (CHAT), educational activities can be described in terms of interrelated elements, i.e., subject, object, tools, rules, division of labor, and community, that influence learning outcomes. A key issue is how these elements mediate students' collaborative efforts and how they contribute to learning. Research Questions: 1) How is collaborative learning implemented in engineering design education? 2) How do the elements of CHAT and their interrelations mediate collaborative learning? and 3) What is the evidence that the implementation of collaborative learning contributed to the achievement of desired learning outcomes? Methodology: A systematic literature review following preferred reporting items for systematic review and meta-analyses protocols guidelines was conducted, including 111 articles published between 2011 and 2021. CHAT was used as analytical framework. Findings: Collaborative learning was implemented in engineering design activities to develop technical as well as nontechnical skills. For the CHAT elements, it was found that establishing a common object, rules for collaboration, and division of labor are essential for effective collaboration and can be enhanced through digital technologies (tools) and support from a community, for example, educators. Finally, results showed that there is evidence that described implementations contribute to learning. However, this evidence needs to be interpreted with care, due to methodological issues in some included articles.

This work forms part of the lecture material for a 1st year BSc course on aerospace design (AE1222-II) at TU-Delft/faculty of Aerospace Engineering. This document deals specifically with space launch vehicle design and presents the general approach and methods used in conceptual space launch vehicle design in the period from 2010 till present (June 2025). The basic method is structured along 7 steps that represent general steps in solving a (design) problem and includes 1) defining the problem, 2) establishment of requirements, 3) setting up design options, 4) analyzing design options, 5) comparing design options, 6) trading design options and selecting the best option choice, and 7) Evaluating and if necessary iterating on the design. In this document all these steps are described in some detail. Analysis methods described allow to study single stage and multi-stage launch vehicles employing a mix of series and parallel staged vehicles, thereby taking into account different rocket propellants. Also, the fundamentals of optimum staging are discussed and worked out. Aspects tackled in the design next to propulsion include aerodynamics and stable flight, rocket structures (including the effect of tank pressure on stage design), avionics, and onboard electric power systems. Additionally, some simple models are introduced that allow for explaining the fundamentals of reliability and cost estimation. For each of the elements covered examples are provided as to help clarify the methods as well as their usage.

To adequately prepare engineering students for their professional career, educational institutions offer projects in which students collaboratively solve engineering design problems. It is known from research these projects can lead to a variety of learning outcomes and student experiences. However, studies that provide insights in the influence of different features of an educational design are rare. In the current study we use Cultural Historical Activity Theory (CHAT) as analytical framework to understand how different elements of an educational design affect students’ experience. Additionally, we use the notion of contradictions to identify opportunities for structural course improvement. Focus groups were conducted with 12 Master students in Aerospace Engineering, that participated in a collaborative engineering design course. During the course, students applied Systems Engineering (SE) and Concurrent Engineering (CE) and worked in the Collaborative Design Laboratory (CDL), which is a state-of-the-art facility that holds a variety of industry relevant tools. It was found that students valued the guidance of their coach and experts, co-located collaboration and the freedom to structure their own process. However, they perceived challenges with regard to adoption of tools in the CDL, sharing their progress with their supervisor, coordination of collaborative efforts and scheduling issues. An analysis using CHAT revealed what contradictions caused these challenges. Finally, recommendations are given on how course structure can be structurally improved.

The space community is currently focusing on defining mission architectures able to perform multiple interplanetary missions to support deep space exploration. In particular, placing orbital propellant depots in strategic locations in space would allow to increase the useful mass transferred. The design of the propellant depot depends greatly on the propellant storage duration and the thermal environment the depot experiences. Furthermore, different cryogenic propellant combinations are being considered for use, including hydrolox and methalox. For both, efficient boil-off reduction strategies are fundamental. The aim of this work is to evaluate different depot architectures for different thermal environments and mission durations. The approach taken in this work included the development of a propellant depot sizing model that allows determining the effect of different thermal control design options, thermal environments, and depot configurations for varying mission duration. The design options include, amongst others, Multi-Layer Insulation and Vapor Cooled Shields. The model also allows for a multi-nodal thermal analysis to estimate boil-off rates for the different designs. Main objective for the studies is to identify the architecture that is most mass efficient. Preliminary results show that mass efficient designs can be achieved with only passive insulation for mission durations below one year, with further improvements when adding a vapor cooled shield to the design.

This study investigates the potential of a newly released multi-phase solver to simulate atomisation in an air-blast type atomiser. The 'VOF-to-DPM' solver was used to simulate primary and secondary atomisation in an atomiser with a coaxial injector-like geometry. The solver uses a hybrid Eulerian/Eulerian-Lagrangian formulation with geometric transition criteria between the two models. In this study isothermal, non-reacting flow at room temperature was assumed. The primary focus was predicting Sauter mean diameter and droplet velocity data at a sampling plane downstream of the injector. The solver produces the expected data and predicts trends similar to those found in experimental measurements. The accuracy of the produced droplet diameters was roughly a factor 2 off compared to experiment. This is attributed primarily to mesh resolution. It was concluded that the solver has the potential to predict atomisation at a reasonable computational cost, but further study is needed to confirm its full capabilities.

A subscale, research rocket thrust chamber operating with cryogenic oxygen and hydrogen exhibits self-excited transverse-mode instabilities with amplitudes of more than 80% of the steady combustion chamber pressure (peak-to-peak) for some operating conditions. During unstable combustion, an increase in the integral heat flux into the water-cooled combustion chamber walls of 20–40% with respect to stable conditions was experienced. A model was derived to predict changes in the axial heat flux profile considering only the dependence of flame length on the amplitude of transverse acoustic oscillations. The model predicts an increase in heat flux in the upstream part of the chamber by up to a factor of 7. This drastic increase is in agreement with past observations of rocket engine failures due to instabilities, in which the structural damage is commonly observed on the faceplate and the walls adjacent to the injection plane. The model also predicts a peak increase in integral heat flux of up to about 25%. While falling short of the peak experimental value of 40%, it nevertheless suggests that flame length is the dominant influence on the distribution of thermal loads in this study.

This work forms part of the lecture material for a 1st year BSc course on aerospace design (AE1222-II) at TU-Delft/faculty of Aerospace Engineering. This document deals specifically with the design of spacecraft and more in particular the bus or the carrier vehicle and presents the general approach and methods used in conceptual spacecraft vehicle design in the period from 2010 till present (June 2025). The basic method is structured along 7 steps that represent general steps in solving a (design) problem and includes 1) problem definition, 2) establishment of requirements, 3) setting up design options, 4) analyzing design options, 5) comparing design options, 6) trading design options and selecting the best option, and 7) evaluating and if necessary iterating on the design. In this document all these steps are described in some detail. Focus is on analysis methods used in conceptual design of a spacecraft with attention to subsystem design. Subsystems dealt with include structures, thermal, propulsion, communications, command and data handling and electrical power.

Since there is a high interest in the use of green propellants, hydrogen peroxide is coming back after once making place for the rise of Hydrazine in monopropellant propulsion systems. Typically, these thrusters are outfitted with catalyst beds. A fully modular 1N thruster is designed to provide the capability of testing and comparing the performance of different concentrations of hydrogen peroxide, different catalysts as well as new technologies in an attempt to resolve the disadvantages associated with the use of catalyst beds. A preliminary baseline design of a catalytic thruster has been created. This will be followed by the design of a secondary decomposition chamber for new technologies, a propellant feed system, a test setup and a test plan.

Delft University of Technology is currently developing the pico-satellite platform Delfi-PQ, based on the PocketQube standard, in pursuit of a new generation of satellites with lower cost, flexibility and short development time. A technology demonstration payload expected to fly in one of the first Delfi-PQ satellites is a dual thruster micro-propulsion system based on the use of water as propellant. Two different micro-resistojet concepts will be demonstrated in the same satellite flight: one based on vaporization, heating and expansion in a nozzle of pressurized liquid water (Vaporizing Liquid Micro-resistojet); the other based on heating and acceleration in slots with simple geometry of molecules of vapour under transitional or free molecular flow regime (Low Pressure Micro-resistojet). The demonstrator is based on a common propellant storage for the two micro-propulsion concepts, based on the use of the capillarity properties of water in a small diameter tube connected to the two separate MEMS thruster chips with their own dedicated valves. This paper describes the requirements and design of the complete micro-propulsion demonstrator as well as its expected operational envelope for in-orbit functional testing, based on the currently validated performance characteristics of the two thrusters.

Commercial launch service providers’ low-priced offerings have been a hotly debated topic. However, the strategies with which these firms reduce costs have seen little incorporation into the hardware Cost Estimating Methods (CEMs) and tools prevalent in the aerospace industry. This research changes this, by providing adaptations to agency-focused CEMs that befit a new commercial paradigm, with an emphasis on smallsat launch vehicles. A parametric model for estimating costs in an early phase of development was synthesized, with which it is possible to approximate the full life-cycle costs of small commercial liquid and solid propellant rockets, as well as their cost-based price per flight. Key elements included were reductions in cost achieved by commercial launch operators, by modeling reduced subcontractor management effort and profit retention experienced at lower subcontracting rates. Prices per flight of small commercial launch vehicles were approximated by combining a parametric cost estimating methodology used frequently in the context of space agencies such as ESA and NASA, called the T1 Equivalents method, with another parametric three-part estimate developed by Koelle for development, manufacture and operations phase costs. The first two phases were estimated through the T1 method, while the operations costs were modeled with TRANSCOST. Along with the newly developed methodologies, novel insights such as required launch rates have shone a light on small commercial launch systems’ cost feasibility in the age of public-private spaceflight partnerships. The model developed was able to approximate costs of development, manufacture and price per flight of three commercial rockets to within 20% of actual reported costs or prices. However, it is recommended the model is refined as more reference cost data, especially on a subsystem level, as well as pricing for these smaller rockets becomes available in the coming years.

Delft University of Technology is currently developing the pico-satellite platform Delfi-PQ, based on the PocketQube standard, in pursuit of a new generation of satellites with lower cost, flexibility and short development time. A technology demonstration payload expected to fly in one of the first Delfi-PQ satellites is a dual thruster micro-propulsion system based on the use of water as propellant. Two different micro-resistojet concepts will be demonstrated in the same flight of the satellite: one based on vaporization, heating and expansion in a nozzle of pressurized liquid water (Vaporising Liquid Micro-resistojet); the other based on heating and acceleration in slots with simple geometry of molecules of vapour under transitional or free molecular flow regime (Low Pressure Micro-resistojet). The demonstrator is based on a common propellant storage for the two micro-propulsion concepts, based on the use of the capillarity properties of water in a small diameter tube connected to the two separate MEMS thruster chips with their own dedicated valves. This paper describes the requirements and design of the complete micro-propulsion demonstrator as well as its expected operational envelope for in-orbit functional testing, based on the currently validated performance characteristics of the two thrusters.

Microelectromechanical systems (MEMS) techniques uncovered new opportunities in satisfying the mission requirements of the growing next generation nano- and pico-satellite missions. In particular, micro-propulsion is universally recognized as one of the key enabling technologies to help this class of satellites making the next step and become credible candidates to a wide range of scientific and commercial applications. In this context, TU Delft is developing a miniaturized electro-thermal propulsion system operating with green liquid propellants, for application on a wide range of nano-satellite formats from CubeSats (10x10x10 cm units) to PocketQubes (5x5x5 cm units). A breadboard of the complete micro-propulsion system is under development at TU Delft, including the thruster, propellant tank, the valve and the driving electronics. The design of the system shall be easily adapted to both CubeSat and PocketQube standards, with particular attention to the second one since the system is scheduled for an initial flight demonstration on the first Delfi-PQ satellite. To address this need and to fill an existing gap in the state-of-the-art of micro-propulsion, two kinds of micro-thrusters are considered in this development; a Vaporizing Liquid Micro-resistojet (VLM) and a Low Pressure Micro-resistojet (LPM). A number of test results will be shown in the paper on the electrical, mechanical and functional characterization of the MEMS thrusters, fabricated in the Else Kooi Laboratory at TU Delft, and the other components of the system. Keywords: Micro-resistojet, Microthruster, MEMS, Cubesats, Pocketqubes

The purpose of this paper is to explain the design process of the pressurised propellant tank for a water micro-resistojet, with application on nanosatellites. The motivation of this propulsion system is its use on future nanosatellite missions of the Delft University of Technology, following the legacy of Delfi C3 and Delfi-n3Xt. The water-based micro-resistojet targets thrust levels in the 0.5-10 mN range by employing MEMS technology in the heating chamber. In order to store the water (propellant), the pressurised tank shall withstand 10 bar as MEOP and have a capacity of 200 ml with a mass limit of 200 grams. The design of the propellant tank is tackled following the Design Trades: a systems engineering approach to the decision-making process. This 10-step methodology begins by framing the decision to be made and outlining the design objectives. Subsequently, the design alternatives are created and analysed via trade-off evaluations, considering the previously-defined objectives. From this analysis, it is possible to draw conclusions that will eventually drive the decision. However, the uncertainty of the trade-off process needs to be assessed, as well as its impact on the final decision. An improved set of design alternatives is developed and new trade-offs are performed, following an iterative process. Finally, an undisputed winning design alternative is achieved and selected, closing the decision process. Regarding the propellant tank, the main objectives of the design are the minimization of cost and risk, with maximum performance and design simplicity. An initial set of characteristics of the design is considered: material, manufacturing process, shape and configuration of the propellant tank. Through table trade-off analyses, the different design alternatives are assessed, and sensitivity analyses are conducted to show the impact of the trade-off weights on the outcome. After a decision on the general characteristics, the lid-tank and tank-bus structure interfaces are evaluated. With the help of Design Option Trees, 5 conceptual designs are modelled in 3D to better understand each design solution. Through an advantages-disadvantages table, the interface characteristics are assessed, and the conceptual designs are refined into a single preliminary design, ready for future structural analysis. In conclusion, a successful implementation of the methodology is achieved. The complete process is documented in order to allow traceability of the decisions taken, facilitating design iterations. Furthermore, a set of recommendations for future work is developed, which include a close relation with the manufacturer from the start and parametrisation of designs.

This paper describes the current developments regarding cryogenic rocket engine technology at Delft Aerospace Rocket Engineering (DARE). DARE is a student society based at Delft University of Technology with the goal of being the first student group in the world to launch a rocket into space. After the launch of the hybrid engine powered Stratos II+ sounding rocket in October 2015, DARE decided to investigate highly efficient liquid rocket engine technology . In this context DARE initiated the cryogenic project with the goal of developing a liquid rocket engine using liquid oxygen and liquid methane as propellants with a nominal thrust in the order of 10kN. Eventually this engine shall power a future sounding rocket into space. As an intermediate step, a 3 kN class engine is being developed. Subsystem development tests and possibly a hot-fire test campaign on this engine are planned for 2016 and its development intends to provide DARE with the required experience and knowledge to develop large scale liquid rocket engines. The engine is developed in cooperation with Heliaq Advanced Engineering and is designed to meet the requirements of the second stage engine of the ALV reusable launch vehicle. The design is a pressure-fed engine, regeneratively cooled using the liquid methane fuel. After passing the coolant channels, the methane is injected into the combustion chamber in gaseous state together with the liquid oxygen in a co-axial manner. The engine is ignited by means of a pyrotechnic igniter using an ammonium perchlorate based propellant. During a test sequence, the propellants are stored in insulated run-tanks and are pressurized using helium. This paper describes the project objective, the current progress on the design and production, and finally four proposed research topics that are indented to be conducted at the faculty of Aerospace Engineering of Delft University of Technology.

In the last decade, CubeSat development has shown the potential to allow for low-risk, low-cost space missions. To further improve the capabilities of CubeSats in large scale missions, a novel micro-propulsion system is being developed at Delft University of Technology. The system is based on a Vaporizing Liquid Microthruster (VLM), which is manufactured by means of Micro Electro-Mechanical Systems (MEMS) technology. It aims to achieve a specific impulse of 100 s and thrust of 1.4 mN, using water as propellant. This paper presents a status update of the development project. Design solutions are shown to circumvent manufacturing tolerances in the wafer-bonding and sealing of the interfaces of the VLM. Secondly, performance analysis based on a 1D-flow approximation is shown to provide a useful tool to quickly predict VLM performance. Next, a detailed design of the propellant storage tank for the CubeSat micro-propulsion system is presented. Finally, the test plan and test setup for the VLM are elaborated, presenting solutions to determine chamber temperature and pressure without directly sensing it.

Methane is a promising propellant for liquid rocket engines. As a regenerative coolant, it would be close to its critical point, complicating cooling analysis. This study encompasses the development and validation of a new, open-source computational fluid dynamics (CFD) method for analysis of methane cooling channels. Validation with experimental data has been carried out, showing an accuracy within 20 K for wall temperature and 10% for pressure drop. It is shown that the turbulence model has only a limited impact on the simulation results and that the wall function approach generates valid results. Finally, a cooling analysis is performed to compare two thrust chamber materials. A traditional copper alloy is compared to aluminium as chamber material for a small moderate-pressure oxygen/ methane engine. The analyses show that aluminium is a feasible chamber material only if a thermal barrier coating is applied. In addition, a significantly higher cooling channel pressure drop is incurred for an aluminium chamber than for a copper chamber due to the lower allowable temperature.