The space sector has been evolving due to the fast-technological advancements generating a reduction of manufacturing, cost and size in space missions, where highly capable performing small satellites are becoming the standard in this industry. Furthermore, the high launching cost limits the trend of cost reduction for the space missions, since the small satellites are sent as a second payload. An alternative to reduce this limitation is using reusable launchers which are key in the future of space industry, once they are optimized in efficiency and reliability. Therefore, an opportunity of design is presented, since the increase of small satellites missions requires a reduction of the cost in launch services a suitable option for the future market are the reusable launchers. The problematic of using recovery systems and reuse parts of the vehicle is the increase of weight due to the added systems that the vehicle needs to be recovered. This paper presents different engines and calculate the performance of each engine based on the needs of missions for small satellites. The starting conditions will be that the payload needs to be launched in low circular or elliptical orbits (altitudes of between 300 and 650 km) and the engine has the ability of vertical take-off, vertical landing. The design will also take into account the possibility of reusing parts of the vehicle and the reentry capability. Different combination of engines and fuels are setup in various configurations. For each case the mass analysis will be developed which will allow to calculate the performance for each engine. The important parameters are the number and type of engines, the ratios of the masses, the thrust-to-weight ratio and specific impulse. Once the mass analysis is obtained the following procedure is the selection of the design considering the empty mass. The best combination of characteristics of the engines will be the suitable candidate. Different assistance systems and techniques for the recovery are assessed to obtain a suitable option to improve the efficiency. The expected results are the calculation of the engine performance and how the selected design can be suitable for the space launcher sector for the small satellites. The expected results are a feasible vehicle for small satellites design based in the calculation of the engine parameters together with an efficient launch recovery system. The conclusion is that the space sector can benefit from the design, demonstrating that a launch vehicle with the reusable characteristics.

Hypersonic propulsion has become an increasingly important research field over the past fifty years, and subsequent interest in propulsion systems utilizing supersonic combustion has emerged. Air-breathing engines are desirable for such applications as hypersonic flight vehicles would not need to carry an oxidizer. Therefore, hypersonic air-breathing propulsion systems require an inlet with high mass capture and compressive efficiency. The present work seeks to outline the development and validation of a novel design tool for producing air inlet designs for hypersonic vehicles at variable flight conditions. A Busemann inlet was chosen for its high compressive efficiency, geometric flexibility, and existing experimental validation. The design tool uses the Taylor-Maccoll equation to generate a streamline through a conical flow field. A streamline tracing technique is used to produce three-dimensional inlet surfaces with various capture areas. Additionally, a surface morphing process is implemented to combine inlet profiles for improved engine compatibility. The inlet morphing process allowed for the creation of inlets with offset exit profiles. These offset profiles were evaluated at off-design Mach numbers using Star-CCM+ to quantify efficiency metrics and characterize starting phenomena.

A facility was designed to provide high-enthalpy, hypersonic flow to a detonation chamber. Preliminary investigation identified 1300 K and Mach 5 as the total temperature and Mach number require to stabilize an oblique detonation wave inside the detonation chamber. Vitiated-air heating was the preheating method chosen to meet these capabilities. The vitiator facility heats compressed air while still retaining about 50% of the original oxygen content. Schlieren flow visualization and conventional photography was performed at the exit plane of a choke plate, which simulated the throat of a converging-diverging nozzle. A shock diamond formation was observed within the jet exhausting out of the choke hole. This is a clear indication that the facility is capable of producing hypersonic flow. A stoichiometric propane-air mixture was burned inside the combustion chamber. A thermocouple survey measured an average temperature of 1099 K at the exit plane of the mixing chamber; however, the actual temperature is likely higher than this, because cool, ambient air could be seen mixing with the hot, vitiated air near the exit plane. Because the adiabatic flame temperature of propane-air is lower than that of hydrogen-air, if hydrogen is used to vitiate the air, the facility is capable of meeting the 1300-K objective.

Trajectory optimization is a powerful technique to analyze mission feasibility during mission design. High-thrust trajectory optimization problems are typically formulated as discrete optimization problems and are numerically well-behaved. Low-thrust systems, on the other hand, operate for significant periods of the mission time. As a result, the solution approach requires continuous optimization; the associated optimal control problems are in general numerically ill-conditioned. In addition, case studies comparing the performance of low-thrust technologies for space travel have not received adequate attention in the literature and are in most instances incomplete. The objective of this dissertation is therefore to design an efficient optimal control algorithm and to apply it to the minimum-time transfer problem of low-thrust spacecraft. We devise a cascaded computational scheme based on numerical and analytical methods. Whereas other conventional optimization packages rely on numerical solution approaches, we employ analytical and semi-analytical techniques such as symmetry and homotopy methods to assist in the solution-finding process. The first objective is to obtain a single optimized trajectory that satisfies some given boundary conditions. The initialization phase for this first trajectory includes a global, stochastic search based on Adaptive Simulated Annealing; the fine tuning of optimization parameters — the local search — is accomplished by Quasi-Newton and Newton methods. Once an optimized trajectory has been obtained, we use system symmetry and homotopy techniques to generate additional optimal control solutions efficiently. We obtain optimal trajectories for several interrelated problem families that are described as Multi-Point Boundary Value Problems. We present and prove two theorems describing system symmetries for solar sail spacecraft and discuss symmetry properties and symmetry breaking for electric spacecraft systems models. We demonstrate how these symmetry properties can be used to significantly simplify the solution-finding process. / Ph. D.

The goal of this design process was to achieve the most efficient propulsive system for the candidate autonomous underwater vehicle (AUV) as possible. A mathematical approach, using fundamental motor equations and derived quantities, was used to characterize and select an efficient brushless electric motor for the propulsion system. A program developed at MIT, Massachusetts Institute of Technology, called OpenProp versions 1 and 2.3 was utilized to design a custom propeller that maximizes the efficiency of the system. A brushless electric motor was selected for the candidate AUV based on a survey of available off the shelf motors and a mathematical characterization process. In parallel with the motor characterization a propeller design was optimized using OpenProp v1 to perform a parametric analysis. OpenProp v2.3 was then used to design a unique propeller for the selected motor. The propeller design resulted in a final propeller with an efficiency of 79.93%. The motor characterization process resulted in two candidate motors being selected, the NeuMotor 1925-3Y and NeuMotor 1521-10.5Y, for in house testing and evaluation. A total propulsive system efficiency of between 44% and 46% was achieved depending on which motor is selected for the final design. / Master of Science

The COLISEUM framework is a tool designed for electric propulsion plume interactions. Virginia Tech has been developing a module for COLISEUM called DRACO, a particle-in-cell based code capable of plume modeling for geometrically complex spacecraft. This work integrates a charging module into DRACO. Charge is collected via particle impingement on the spacecraft surface and converted to potential. Charge can be stored in the surface, or added to a local ground potential. Current can flow through the surface and is governed by the internal electric field in the spacecraft. Several test cases were run to demonstrate the code's capabilities. The first was a plume impingement of a composite spherical probe by a xenon thruster. It was shown that the majority of current conducted will reach the interior of the spacecraft, not other surface elements. A conductive interior will therefore result in a uniform surface potential, even for low surface conductivities. A second test case showed a composite spacecraft exposed to a solar wind. This test showed that when a potential gradient is applied to a conductive body, the ground potential of the spacecraft will lower significantly to compensate and maintain a zero net current. The case that used the semiconductive material showed that the effect of the potential gradient can be lowered in cases such as this, where some charge will always be stuck in the surface. If a dielectric material is used, the gradient will disappear altogether. The final test case showed the effect of charge exchange ion backflow on the potential of a spacecraft similar to the DAWN spacecraft. This case showed that CEX ion distribution is not even along the spacecraft and will result in a transverse potential gradient along the panel. / Master of Science

A trade study is carried out for the design of electric propulsion based lunar robotic precursor missions. The focus is to understand the relationships between payload mass delivered, electric propulsion power, and trip time. The results are compared against a baseline system using chemical propulsion with LOX/H2. The major differences between the chemical propulsion based and electric propulsion based systems are presented in terms of the payload mass and trip time. It is shown that solar electric propulsion offers significant advantage over chemical propulsion in delivering non-time critical payloads to lunar orbit. / Master of Science

Underwater surveillance missions both for defense and civilian applications are continually demanding the need for unmanned underwater vehicles or UUVs. Unmanned vehicles are needed to meet the logistical requirements for operation over long distances, greater depths, long duration, and harsh conditions. In order to design UUVs that not only satisfy these needs but are also adaptive and efficient, there has been increasing interest in taking inspiration from nature. These biomimetic/bio-inspired UUVs are expected to provide significant improvement over the conventional propeller based vehicles by taking advantage of flexible bodies and smart actuation. In this thesis, jetting jellyfish were utilized as the inspiration to understand the fundamentals of this new form of propulsion and subsequently translate the understanding onto the engineered platform to validate the hypothesis and construct robust models. Jetting jellyfish species are generally smaller in dimensions than rowing jellyfish, consume lower energy for transport, and exhibit higher proficiency. In the second chapter, a bio-inspired stationary jet propulsion mechanism that utilizes an iris diaphragm actuation system was developed. Detailed discussion is provided on the design methodology and factors playing the leading role in controlling the vortex formation. The propulsion mechanism was intended to mimic the morphological and deformation features of Sarsia sp. jellyfish that measures approximately 1 cm in diameter. The performance of experimental model was analyzed and modeled to elucidate the role of structure and fluid displacement. Utilizing the results from Chapter 2, a free-swimming jellyfish-inspired robot (named JetPRo) was developed (also utilizing an iris diaphragm) in Chapter 3 and characterized for relevant propulsive metrics. A combination of theoretical modeling and experimental analysis was used to optimize the JetPRo's gait for maximum steady-state swimming velocity. Next, an attempt was made towards creating a free-swimming jetting robot (named JP2) using a guided cable mechanism to achieve the desired actuation and improve the propulsion while simplifying the drive mechanism. Using JP2 robotic model, a systematic set of experiments were conducted and the results were used to refine the theory. Based upon the comprehensive computational analysis, an optimized swimming gait was predicted and then validated. A modular robot inspired by siphonophores was developed and initial efforts were made in laying down the foundation for understanding of this complex locomotion mechanism. Siphonophores are colonial organisms consisting of several jetting bodies attached to a central stem. An experimental model was developed mimicking the multimodal swimming propulsion utilized by Siphonophores. Several swimming gaits inspired by the natural animal were replicated and the preliminary performance of the experimental model was quantified. Using these results, an analysis is presented towards further improving the design and assembly of a siphonophore-inspired robot. / Master of Science

A first-generation LaB6 heaterless hollow cathode with a flat-plate anode is experimentally investigated. The cathode is characterized using krypton as propellant at varying flow rates, discharge currents and cathode-anode distances. Voltage probes, used to make direct voltage measurements in the ignition circuit, are the only diagnostic tool used experimentally. A plasma model is used to infer plasma parameters in the cathode emitter region. The cathode characterization results are consistent with those obtained during previous investigations of 1 A-class LaB6 hollow cathode with krypton. A peak-to-peak anode voltage criterion is used to identify the discharge modes and the occurrence of mode transition. Fourier analysis of the keeper and anode voltage waveforms carried out to study the discharge mode behavior reveals resonant frequencies ranging from 40 to 150 kHz. Lastly, post-test visual observations of the cathode components show signs of emitter poisoning and keeper erosion. / Master of Science / Recent years have seen rapid growth in the development of both stand-alone satellites and satellite constellations. A critical component of these satellites is the on-board propulsion system, which is responsible for controlling their orientation with respect to the object of interest and keeping the spacecraft in the assigned orbit. Generally, electric propulsion systems are used for this purpose. These types of propulsion systems use electrical power to change the velocity of satellite, providing a small thrust for a long duration of time as compared to chemical propulsion systems. Certain types of electric thrusters utilize a hollow cathode device as an electron source to start-off and support the thruster operation. In this research, a non-conventional hollow cathode for low power applications is developed and tested. The main characteristic of the developed cathode is the heaterless configuration, which eliminates the heater module used in conventional cathodes to enable the cathode to reach its operational temperature. The absence of a heater reduces the complexity of the cathode and the electrical power system. The cathode utilizes an electron emitter material which is insensitive to impurities and air exposure. Additionally, unlike typical electric thrusters which use xenon as the fuel, this cathode uses krypton which is similar to xenon but is less expensive. The presented work includes an overview of electric propulsion and the hollow cathode operation, followed by a detailed discussion of the heaterless hollow cathode design, the experimental setup and the test results. Several noteworthy findings regarding cathode operation are included as well. This research shows that the non-conventional heaterless hollow cathode and its operation with krypton have the potential to improve the overall thruster performance by reducing the weight and the cost, thus contributing to an integral aspect of satellite on-board propulsion.

A direct connect, supersonic solid fuel combustor with a cavity is explored in the context of understanding characteristics related to ignition, regression rate, combustion, and flow fields for application in advancing solid fuel scramjet research. 3D printed, polymethylmethacrylate fuel grains are loaded into both fully enclosed and optically accessible combustors. The ignition characteristics are investigated by systematically varying the internal geometry of the fuel grain to develop a flammability map with respect to non-dimensional geometric parameters. Results reveal that a longer and larger flameholding cavity creates favorable conditions for ignition and sustained combustion. The inlet temperature is also systematically varied to extend the available literature on the supersonic combustion of solid fuels to lower temperature operating conditions and show that a higher inlet temperature is conducive to sustained combustion and higher regression rates. The regression rates of the fuel grains are measured to determine a concentration of regression in the flameholding cavity along the angle of the downstream side of the cavity. Ignition and sustained combustion rely heavily on the fuel in the flameholding cavity. A decreasing regression rate is observed as the fuel regresses by measuring the regression rate at discrete time intervals during a firing of the optical combustor. The optical combustor is also subject to various high-frequency imaging techniques. Shadowgraph imaging shows the changes in density of the flow field and finds a normal shock in the constant area section. CH* chemiluminescence imaging provides novel observations of the concentrated areas of combustion along the fuel grain wall by highlighting the heat release from combustion. A high intensity of CH* radicals is in the upstream section of the flameholding cavity. When considered in the context of the concentration of regression, this indicates that the recirculation zone pulls fuel from the downstream section of the cavity, combusts it in the upstream section of the flameholding cavity, then expels the higher enthalpy gas into the core flow. Additionally, observing the flow provides insight into the flow dynamics of opposing cavities in a supersonic flow field. The symmetry of the flow field is found to be reliant on the stability of the flameholding cavity length to depth ratio. / Master of Science / A solid fuel scramjet has the potential to be the simplest and most cost effective method of achieving hypersonic flight. A liquid fuel scramjet has been demonstrated in free flight, but liquid fuels present many issues involving safety and storage that can be eliminated by introducing solid fuels. Supersonic combustion, or burning fuel in an air flow moving faster than the speed of sound, is a complicated subject due to the irregularity of flow fields and the requirement of combustion to occur at a high rate. The research within this thesis presents many novel technologies that have never been presented in published literature in the context of the supersonic combustion of solid fuels. By conducting ground testing of a solid fuel scramjet, characteristics of the combustion can be studied to expand the available literature in the field to new fuel geometries and inlet conditions. The ignition and sustained combustion of a solid fuel scramjet is extremely reliant on the initial geometry of the fuel and the initial temperature of the flow. This research advances the field of supersonic combustion of solid fuels by developing an optically accessible combustor using quartz windows. These characteristics of supersonic combustion are investigated using highspeed video recording. The results of these techniques provide insight into favorable fuel geometries and inlet conditions. Additionally, patterns observed in the flow field explain concentrations of combustion and fuel consumption.