The Auroral Spatial Structures Probe (ASSP) mission is a sounding rocket mission studying solar energy input to space weather. ASSP requires the high velocity ejection (up to 50 m/s) of 6 secondary payloads, spin stabilized perpendicular to the ejection velocity. The proposed scientific instrumentation depends on a high degree of spin stability, requiring a maximum coning angle of less than 5º. It also requires that the spin axis be aligned within 25º of the local magnetic field lines. The maximum velocities of current ejection methods are typically less than 10m/s, and often produce coning angles in excess of 20º. Because of this they do not meet the ASSP mission requirements. To meet these requirements a new ejection method is being developed by NASA Wallops Flight Facility. Success of the technique in meeting coning angle and B-field alignment requirements is evaluated herein by modeling secondary payload dynamic behavior using a 6-DOF dynamic simulation employing state space integration written in MATLAB. Simulation results showed that secondary payload mass balancing is the most important factor in meeting stability requirements. Secondary mass payload properties will be measured using an inverted torsion pendulum. If moment of inertia measurement errors can be reduced to 0.5%, it is possible to achieve mean coning and B-field alignment angles of 2.16º and 2.71º, respectively.

The aim of this thesis is to look at the possibility for SSC, Swedish Space Corporation, to build a cost-effective mobile rocket launcher to be used for their most common sounding rockets. Having a mobile rocket launcher will give SSC the possibility to not only expand their own line of launchers but more importantly giving them the opportunity to launch rockets outside their own base. To ease transportation of the launcher it is required to fit in a 40ft container. This requirement is the major limitation during the design phase. To keep the cost down, emphasis will be put on trying to find solutions with commercial products. Concepts were developed during three phases and after presenting the concepts to involved personnel, feedback was given and the concept(s) were developed further. The third and final phase contain the chosen concept and suggestions are given regarding future work to be done before it can be manufactured.

Due to renewed interest in the sub orbital rocket program of the Florida Space Authority and a surplus of Super Loki Sounding Rockets, an effort to improve the usefulness of this surplus is herein undertaken. Currently, the capacity of the payload compartment in the upper stage of the Super Loki system is very limited. A redesign of the upper stage will allow larger and more versatile payloads to be carried, assuming the appropriate design parameters are met. It has therefore been undertaken to create a design procedure that is comprehensive in scope in order to affect this redesign. This procedure includes five major components. These are the separation of the upper and lower stages, the stability of the vehicle, the altitude and velocity of the rocket, the mechanical loading and finally the aerodynamic heating. Semi-empirical methods were used whenever possible to allow comparison with experimental data. This procedure revealed that larger diameter upper stages might be used up to a reasonable maximum of four inches. The four-inch modification is found to be stable as were the smaller modifications considered. The altitude and velocity of the rocket were found via a simple Eulerian time stepping scheme resulting in an estimate of approximately 148,000ft for the four-inch dart. The mechanical loading analysis allowed for the material selection for the rocket components. Reinforced steel fins and carbon fiber tubing, for the payload section, are adequate to meet expected mechanical loads, those being, 16000psi for the fin section due to launcher forces, 22800psi for compressive plus torsion forces on the composite section and 18000psi for the ejection stresses. An ablative coating is considered necessary to counteract the 760ºF temperatures along the composite tube.

abstract: The development of new Ultra-Violet/Visible/IR range (UV/Vis/IR) astronomical instrumentation that use novel approaches for imaging and increase the accessibility of observing time for more research groups is essential for rapid innovation within the community. Unique focal planes that are rapid-prototyped, low cost, and provide high resolution are key. In this dissertation the emergent designs of three unique focal planes are discussed. These focal planes were each designed for a different astronomical platform: suborbital balloon, suborbital rocket, and ground-based observatory. The balloon-based payload is a hexapod-actuated focal plane that uses tip-tilt motion to increase angular resolution through the removal of jitter – known as the HExapod Resolution-Enhancement SYstem (HERESY), the suborbital rocket imaging payload is a Jet Propulsion Laboratory (JPL) delta-doped charge-coupled device (CCD) packaged to survive the rigors of launch and image far-ultra-violet (FUV) spectra, and the ground-based observatory payload is a star centroid tracking modification to the balloon version of HERESY for the tip-tilt correction of atmospheric turbulence. The design, construction, verification, and validation of each focal plane payload is discussed in detail. For HERESY’s balloon implementation, pointing error data from the Stratospheric Terahertz Observatory (STO) Antarctic balloon mission was used to form an experimental lab test setup to demonstrate the hexapod can eliminate jitter in flight-like conditions. For the suborbital rocket focal plane, a harsh set of unit-level tests to ensure the payload could survive launch and space conditions, as well as the characterization and optimization of the JPL detector, are detailed. Finally, a modification of co-mounting a fast-read detector to the HERESY focal plane, for use on ground-based observatories, intended to reduce atmospherically induced tip-tilt error through the centroid tracking of bright natural guidestars, is described. / Dissertation/Thesis / Doctoral Dissertation Exploration Systems Design 2019

<p>This report handles the preliminary design of a control system that includes both attitude control and boost control functionality for sounding rockets. This is done to reduce the weight and volume for the control system. </p><p>A sounding rocket is a small rocket compared to a satellite launcher. It is used to launch payloads into suborbital trajectories. The payload consists of scientific experiments, for example micro-gravity experiments and astronomic observations. The boost guidance system controls the sounding rocket during the launch phase. This is done to minimize the impact dispersion. The attitude control system controls the payload during the experiment phase. </p><p>The system that is developed in this report is based on the DS19 boost guidance system from Saab Ericsson Space AB. The new system is designed by extending DS19 with software and hardware. The new system is therefore named DS19+. Hardware wise a study of the mechanical and electrical interfaces and also of the system budgets for gas, mass and power for the system are done to determine the feasibility for the combined system. </p><p>Further a preliminary design of the control software is done. The design has been implemented as pseudo code in MATLAB for testing and simulations. A simulation model for the sounding rocket andits surroundings during the experiment phase has also been designed and implemented in MATLAB. The tests and simulations that have been performed show that the code is suitable for implementation in the real system.</p>

This report handles the preliminary design of a control system that includes both attitude control and boost control functionality for sounding rockets. This is done to reduce the weight and volume for the control system. A sounding rocket is a small rocket compared to a satellite launcher. It is used to launch payloads into suborbital trajectories. The payload consists of scientific experiments, for example micro-gravity experiments and astronomic observations. The boost guidance system controls the sounding rocket during the launch phase. This is done to minimize the impact dispersion. The attitude control system controls the payload during the experiment phase. The system that is developed in this report is based on the DS19 boost guidance system from Saab Ericsson Space AB. The new system is designed by extending DS19 with software and hardware. The new system is therefore named DS19+. Hardware wise a study of the mechanical and electrical interfaces and also of the system budgets for gas, mass and power for the system are done to determine the feasibility for the combined system. Further a preliminary design of the control software is done. The design has been implemented as pseudo code in MATLAB for testing and simulations. A simulation model for the sounding rocket andits surroundings during the experiment phase has also been designed and implemented in MATLAB. The tests and simulations that have been performed show that the code is suitable for implementation in the real system.

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This thesis has been performed in collaboration with the Swedish Space Corporation at the department Science Services. SSC provides services in the areas of spacecraft subsystems, ground stations and sounding rockets to enable governments, companies and research institutes to benefit from space. Science Services are responsible for sounding rocket flight missions allowing customers to perform research in a microgravity environment. Currently, they have good knowledge how to design the sounding rockets experiment modules to minimize thermal effects within the system. However, no computational models are available to evaluate and verify the thermal heat transfer inside of the modules and as such the systems are designed primarily based on previous experience. The main purpose of this thesis was to develop a thermal computational model, which would work as a basis for designing experiment modules. The model would be used in an early stage of the design process before CAD parts have been designed. This required a flexible model allowing the user to evaluate different types of components and configurations. A finite element method (FEM) was used to perform heat transfer calculations in MATLAB. The development process was divided into three stages, which reduced the complexity of the problem formulation. The first version was made to approximate heat transfer solution in three dimensions using the Galerkin’s weighed residuals method. The second version was made to implement the dynamic environment occurring during flight missions. Based on the external environment, the dynamic process was divided into phases with different boundary conditions. In the final version internal convection, conductivity between air elements and a GUI was developed. The versions were verified with COMSOL (2013) and previous measured flight data. The results from the simulations showed that the internal convection coefficient and the element’s conductivity have a great impact on how the heat is distributed inside th e modules. A low convection will lead to internal temperature peaks, which can cause damage to sensitive experiment equipment. Also, the results indicated that the external environment does not have a significant impact on the internal temperatures. The assumptions made and recommendations are also covered in this thesis. Keywords: Three-dimensional heat transfer, Finite element method, Sounding rocket, Computational simulation / Detta examensarbete har utförts i samarbete med Swedish Space Corporation på avdelningen Science Services. SSC är ett svenskt företag verksam inom rymdtekniksektorn som erbjuder myndigheter, företag och forskarlag runt om i världen möjlighet att dra nytta av rymden. Avdelningen Science Service är ansvariga för utvecklig samt uppskjutning av sondraketer. I dagsläget finns en god kunskap hur sondraketens experimentmoduler ska konstrueras för att minimera den termiska påverkan i systemen. Dock existerar ingen beräkningsmodell för att undersöka värmeutvecklingen och temperaturer i dessa experimentmoduler, all kunskap inom detta område är baserad på tidigare erfarenheter. Syftet med detta examensarbete var att utveckla en termisk beräkningsmodell som kan användas som underlag när nya experimentmoduler konstrueras på SSC. Användningsområdet för modellen var avsett i ett tidigt skede i produktutvecklingsprocessen, innan CAD-modeller eller dylikt har framställts. Därav efterfrågades en flexibel modell där användaren kan undersöka olika typer av kompententer och konfigurationer. Den Finita elementmetoden (FEM) har används för att skapa en termisk beräkningsmodell i MATLAB. Utvecklingen delades upp i tre steg, eller tre programversioner, vilket bidrog till att frågeställningens komplexitet reducerades. Första programversion genomfördes för att approximera värmeflöden och temperaturer i tre dimensioner med hjälp av Galerkins viktade residualmetod. I den andra programversionen implementerades den dynamiska omgivningen som uppstår under flygning. Baserat på den yttre påverkan från det dynamiska förloppet delades flygningen in i olika faser, alla med skilda randvillkor. I den slutliga programversionen implementerades intern konvektion, strålning och ett grafiskt användargränssnitt. Samtliga versioner verifierades numerisk med hjälp av COMSOL (2013) . Resultatet från beräkningsmodellen påvisade att den interna konvektionskoefficient samt konduktiviteten hos element har stor inverkan på hur temperaturen fördelas inuti modulen. Resultaten indikerade även att den yttre miljön inte har en signifikant betydelse för dessa temperaturer. De antaganden som utförts samt förbättringsförslag avhandlades även i detta arbete. Nyckelord: Värmeledning, tredimensionellt, Finita elementmetoden, Sondraket, Simulering

The Kentucky Re-entry Universal Payload System (KRUPS) is an adaptable testbed for atmosphere entry science experiments, with an initial application to thermal protection systems (TPS). Because of the uniqueness of atmospheric entry conditions that ground testing is unable to replicate, scientists principally rely on numerical models for predicting entry conditions. The KRUPS spacecraft, developed at the University of Kentucky, provides an inexpensive means of obtaining validation data to verify and improve these models. To increase the technology readiness level (TRL) of the spacecraft, two sub-orbital missions were developed. The first mission, KUDOS, launched August 13th, 2017 on a Terrier-Improved Malamute rocket to an altitude of ~150 km. The second mission, KOREVET, launched on March 25th, 2018 on the same type of rocket to an altitude of ~170 km. The chief purpose of both missions was to validate the spacecraft design, ejection mechanism, on-board power, data transmission, and data collection. After both missions, the overall TRL improved from 4 to 5 by validating most subsystems in a relevant environment. Both of these missions were invaluable preparation for the project's ultimate goal of releasing multiple experimental testbeds from the ISS.