CubeSat missions have been deployed to cislunar space and beyond, paving the way for the next significant advancement: a dedicated CubeSat mission to explore a planet near Earth. To contribute to this goal, the German Aerospace Center (DLR) has developed radiation-hardened small satellite technologies, including communications, power and onboard computer subsystems. This study presents a stand-alone Mars exploration mission using the CubeSat standard that will demonstrate DLR’s in-house technologies. This mission will perform an independent transfer to the Red Planet, achieve orbital insertion, and conduct measurements on its lower atmosphere and gravity field. A system concept has been created by integrating the in-house technologies, investigating the necessary Commercial-Off-The-Shelf (COTS) components, and performing the mission analysis to assess the feasibility of the mission. The resulting 12U CubeSat has a 20.8 kg wet mass, 6.3 km/s low-thrust maneuvering capability and can generate up to 90 W of power at Mars. The proof-of-concept mission is planned for a 4-year duration.

Observing that the majority of SSTO programs failed due to the propulsion system being far ahead of its time. Furthermore, acknowledging the that the private launch markets is increasingly dominated by partly reusable TSTOs. This thesis aims to evaluate whether next generation rocket engines make an SSTO viable within a foreseeable economical time frame, for the increasingly privatised space industry. From the market analysis it follows that the VT SSTO RLV developed by the private launch industry, with a payload capability of 15- 20 tons to a LEO with an altitude of 200-400 km, is most likely. The evaluation is performed by a two level trade-off. Initially, a literature trade-off is preformed on elaborate range of potential next generation rocket engine, which are subdivided into pure rocket engine and breathing engine. All engines are evaluated on the market requirements, performance, and achievability. Only three engine types are selected. Upon evaluating the propulsion literature it was noticed that no literature cross evaluated advanced propulsion systems with high performance propellant pallets. Therefore, a literature propellant trade-off is performed. For the aerospike, the pulse detonation engine and the precooled hybrid airbreathing rocket engine with the propellant pallets H2/O2, CH4/O2, C2H2/O2, and C2H4/O2 a performance analysis is performed. The performance analysis is done by both optimising the engine configurations and the ascent trajectory for a 15 ton payload delivery to an orbit of 400 km altitude. To simulate each engine, three performance analysis are developed, which are all derived from the continuity equations to ensure a consistent comparison. The aerospike is simulated as a continuous optimal convectional via the frozen equilibrium method, although it is found an alternation is needed to account for directional losses. The open-end of the PDE is simulated via a CJ-detonation wave followed by an exponential decaying blowdown phase. The DC nozzle expansion is iteratively simulated via the frozen equilibrium method. The intake and HPC of the precooled hyrbid airbreathing engine is simulated with convectional aero engine performance analysis, while the precooler is simulated ass a counter-flow heat exchanger. The CC and CD nozzle are simulated via the frozen equilibrium method. From the optimised trajectory the representative performance of each engine is extracted and five engine configurations are identified as viable VT SSTO RLV engines. These are the H2/O2 powered aerospike, the pulse detonation engine and the precooled hybrid airbreathing rocket engine operating on an H2/O2 pallet, and the C2H2/O2 powered aerospike and pulse detonation engine. The H2/O2 powered PDE is found to be the most promising, as it offers the best performance gain relative to its achievability. Therefore, it advocates that future VT SSTO RLV research utilises the H2/O2 powered pulse detonation engine as the main propulsion system

Low-thrust propulsion's main advantage is its efficient propellant usage. This holds for manoeuvres as well as for major orbital changes, both for LEO/MEO/GEO satellites and for interplanetary missions. In designing such missions, extensive numerical models as well as less demanding analytic first-order approximations can be used. A specific kind of the latter category of methods are called shape-based methods. TU Delft's Astrodynamic Toolbox (TUDAT) contains the framework for shape-based methods with spherical shaping and hodographic shaping already implemented for use by students and staff. Having more shape-based method available allows for a more versatile toolbox and combining several shapes with for example coasting would lead to more attractive and efficient transfers. This thesis provides the theory required for the implementation of both exponential sinusoids and the sixth order inverse polynomials with 3D capabilities. These methods have been successfully implemented, verified and validated into the TUDAT shape-based framework.
The four shape-based methods have been applied to an Earth-Tempel-1 transfer for start dates from 2020 to 2025 and time of flights of 100 to 8900 days. It is found that the following order can be established with respect to their lowest ΔV, based on both an unoptimised grid search and global optimisation: Inverse polynomial shaping (10.51 km/s), Spherical shaping (11.71 km/s), Exponential shaping (13.09 km/s), Hodographic shaping (26.22 km/s).
Optimising for ΔV and peak thrust acceleration with the Earth-Tempel-1 transfer including a coasting arc between two powered arcs for inverse polynomials and spherical shaping compared to the standard transfer, yields the following results: Trajectories with similar and lower ΔVs, similar and lower peak thrust for the same ΔV and similar and longer time of flights. In conclusion coasting arcs provide more flexibility in terms of start date, time of flight and peak acceleration thrust for a marginal increase or decrease in ΔV.

For a human Mars exploration mission, it is required to minimize the time-of-flight of the interplanetary trajectories to mitigate the adverse effects of the radiation environment and prolonged weightlessness conditions on the health of the crew. Moreover, having a fail-safe provision to return the crew back to Earth in case of a mission abort situation is highly desirable. This thesis research investigated several optimal/sub-optimal and feasible solutions for the high-thrust interplanetary transfer and abort trajectories of a reconnaissance human Mars mission. A conjunction-class mission architecture has been analyzed to transport the crew to and from Mars during the transfer opportunities in 2028 and 2030, respectively. Baseline requirements of such a mission were defined by finding the optimal ballistic Earth-to-Mars and Mars-to-Earth transfer trajectory solutions. A number of propulsive abort trajectory solutions were then computed that can return the crew back to Earth when the nominal mission is aborted either during the Earth-to-Mars transfer or during the Mars surface stay period. A semi-analytic trajectory model was used for the design of such abort trajectories that can include one or two deep space maneuvers and a powered Mars swing-by. With the use of a meta-heuristic global optimization algorithm, multi-objective optimization problems were solved to minimize the ΔV cost and the interplanetary transit duration of the mission. By comparing the transfer and abort trajectory solutions, it was concluded that various abort trajectory options can be provided for a reconnaissance human Mars mission, that also satisfy the imposed constraint on the arrival excess velocity. The effects of such abort trajectory options on the ΔV cost and other baseline mission requirements (such as the total mission duration) were analyzed. Critical challenges of such a safe human Mars mission architecture have been identified and discussed.

Micropropulsion is universally considered to be a key technology enabling nano- and pico- satellites to perform more complex missions. However, past research has shown that nozzle efficiencies at the microscale are far inferior to their macro scale counterparts. These low efficiencies can be attributed to the relatively high viscous losses associated with this microscale. This thesis conducted a three-dimensional numerical study to investigate the impact of the nozzle geometry on the viscous losses. Furthermore, the designed micronozzles are fabricated and the ground work is laid for experimental testing of these nozzles. Results of the numerical study show that by application of a new double depth micro aerospike design nozzle efficiencies can be improved by as much as 41.2%.

The Deployable Space Telescope (DST) project aims at producing a competitive telescope by using segmented deployable light weighted optics which reduces launch volume and mass, and thus launch costs. Due to the criticality of heat management to the performance of the telescope, a thermal model had to be built in order to determine the expected temperatures throughout the telescope in orbit. Further, the premature baffle design required a supportive thermal analysis such that its critical parameters could be determined.

The graduation project was conducted at the upper stage liquid propulsion department of ArianeGroup at the facilities of Airbus DS, Bremen. Based on a series of past flights of Ariane 5 launcher, the aim is to analyze the fluid motion and the pressure fluctuations in the cryogenic propulsion upper stage Liquid Hydrogen (LH₂) and Liquid Oxygen (LO_x) tanks during the ascent phase. Pressure fluctuations (drops or rises) during the ascent phase are undesirable due to the need for relief or re-pressurization of the fuel tanks. Tank relief is obtained through relief valves, while re-pressurization is done using on-board gaseous Helium; both cases increase the failure probability of the system and/or the total weight of the launcher.The objective of the project is to find out why these pressure fluctuations occur, what are the parameters that affect the pressure evolution and at what extend the liquid fuel motion (sloshing) is responsible for this behavior.According to the literature several parameters affect the pressure evolution during sloshing. These parameters are further investigated through flight data analysis. The approach also involves CFD simulations of the kinematic behavior of the liquid fuel focusing on the sloshing angle. Finally, a statistical model is built attempting to predict the pressure change inside the tanks. Higher sloshing angles match with higher pressure rise inside the LO_x tank. The magnitude of the pressure rise appears to be directly connected to the kinematic profile of the launcher as well as to the ullage volume of the tank. The maximum predicted ullage pressure is below the tank's sizing pressure limit. Regarding the LH₂ tank, no strong correlation of flight parameters to the pressure change is identified; no sufficient statistical model is built. The CFD simulation shows that relatively higher sloshing angle magnitude and duration exists near the pressure drop periods and that strong breaking waves are likely to be formed in the case of a sudden pressure drop behavior. The LH₂ tank is more prone to the formation of breaking/splashing waves. The effect of vibrations, which is not included in the CFD study, is also important for the explanation of the pressure drop magnitude.

Comets, the sporadic visitors from the outer edges of the Solar System, are considered to hold the key for understanding the formation of planets and the origin of life on Earth. Having spent the majority of time away from the radiative environment of the inner Solar System, the chemistry of the comets has remained unaltered, making them the pristine samples of the matter from the ancient Solar nebula. A mission to bring cometary particles back to Earth enables the examination of the materials in well equipped laboratories and saves the mass of the instruments to be carried on board. As conventional propulsion methods require a large quantity of propellant for this type of mission, the feasibility of using the novel propulsion technique of solar sailing is explored in this thesis. In order to return the comet samples to Earth within a reasonable time period, the orbit transfer is considered as an optimal control problem with constraints placed on the sailcraft’s position and velocity. The Differential Evolution (DE) algorithm was used to search for time-optimal trajectories that minimize the approach distance and the relative velocity with respect to the comet during sample collection. The optimal trajectory obtained predicts the solar sailcraft to reach the comet, collect the samples and return back to Earth in 6.8 years. The time of arrival at the comet was found to match with the comet's perihelion passage, enabling effective sample collection. The outcome of the trajectory analysis, thus successfully demonstrates the applicability of solar sailing to comet sample return missions in the near future.

A buckling analysis and imperfection sensitivity study of scaled launch-vehicle cylindrical shells have been executed. The emphasis is on scaled cylindrical shells due to the expensive nature of full scale launch-vehicle cylindrical shells and the size constraints of experimental testing equipment. The work as presented in this thesis is part of a framework of a collaboration with NASA Langley. The main objective was to investigate if a scaling method, which is developed as part of this afore mentioned collaboration, results in representative scaled cylindrical composite shells. These scaled cylindrical composite shells will be validated by experimental tests at NASA Langley. The scaling of structures can be a challenging process, as a scaled model which shows full scale representative behaviour is difficult to design due to constraints such as manufacturability. The buckling of cylindrical shells is considered to be a structural problem which is yet to be fully understood. The cylindrical shells show a high imperfection sensitivity, but the exact influence of imperfections caused by different manufacturing processes is a source of uncertainty. The thesis will focus on two scaled cylindrical shells which resulted from the scaling method, next to the full scale CTA 8.1 cylindrical shell they are based on. The CTA 8.1 is a sandwich cylindrical shell with a honeycomb core which is designed to be representative for a launch-vehicle structure. The first scaled cylindrical shell consists of a solid laminate and the second scaled cylindrical shell consists of a sandwich with foam core. A scaled solid laminate cylindrical shell is of interest, as the core thickness of a scaled sandwich cylindrical shell can become too thin to manufacture...

Over the past years, numerous missions for spacecraft with low-thrust propulsion have been planned to include large orbital plane changes. In order efficiently generate and evaluate orbits for such missions, use can be made of the spherical shaping method proposed by Novak. When inclinations larger than 15 degrees are present however, the error in the ΔV found by this method increases significantly. It was found that the spherical shaping method’s applicability and accuracy at high orbital inclinations can be improved through the development of a more accurate elevation shaping function. The purpose of this MSc thesis was therefore to do research on the development and implementation of such a function.

Using a number of test cases, three potential methods to obtaining a more accurate elevation shaping function were evaluated. These methods included the usage of spherical triangles, Fourier series and an alternative function that was obtained during the development. By evaluating the ability of various functions to describe the unperturbed and perturbed orbits included in the test cases, a new elevation shaping function was found. This shaping function was based on the aforementioned alternative function obtained during the development.

The accuracy of the new elevation shaping function was tested using a number of cases. It was observed that the new elevation shaping function significantly increased the spherical shaping method’s accuracy at high orbital inclinations; at an inclination of 20 degrees, the error in the ΔV found was decreased to less than 1 percent. However, this error did increase as the inclination became larger.

The spherical shaping method was furthermore applied together with the new elevation shaping function to the design of two missions. These missions included rendezvous trajectories to the dwarf planet Makemake and to the comet 2003 EH1. It was found that the new elevation shaping function enabled the spherical shaping method to produce smooth trajectories with attainable ΔVs to both targets, even though the inclination of 2003 EH1 is approximately 70 degrees.