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The Lower Atmosphere/Ionosphere Coupling Experiment (LAICE) CubeSat mission includes the payload instruments that generate scientific data by interacting with the flight computer. The LAICE Instrument Interface Board (LIIB) is designed to interface with the payload instruments and the flight computer for efficient operation of the LAICE. The uplink command packet contains commands for regulating power supply to the payload instruments and for interfacing the peripheral, called the thermal knife, with the science instruments. The LIIB is responsible for interpreting these commands in order to execute the associated functions. The architecture of the LIIB is designed such that it not only takes into account all the requirements of the systems and instruments on the LAICE, but also ensures smooth flight data analysis at the ground station end. The approach taken to build the design makes the entire process intuitive and easier to debug. This thesis describes the design and development of the LIIB firmware, to ensure proper functioning of the LAICE. The firmware design is presented first, by initially defining the architecture based on the system requirements and progressing eventually to its development at the system level. End-to-end testing with the payload instruments and thermal knife setup verifies the operation of the LAICE LIIB firmware and electronics, thus qualifying the instrument for deployment within the LAICE. / Master of Science

For a successful CubeSat mission, it is imperative to schedule events in a fashion that will generate maximum useful science data. Intuitive uplink commanding software is required for the Lower Atmosphere/Ionosphere Coupling Experiment (LAICE) CubeSat to ensure best results. The ground station up-link software is created with this aim in mind. This will make the operation center for the LAICE project efficient. This will also help in evaluating the effect of a particular schedule on LAICE instrument interface board (LIIB) before sending the commands to it. The interactive User Interface (UI) that makes the entire process intuitive guides the user to create an uplink schedule without any human error. The control software creates the command sequence taking in to account all the limitations and specification of the systems and instruments on LAICE. These data are backed up in an efficient format in Virginia Tech’s database for future processing. This web-based application ensures a smooth scheduling process without any errors. Assistive flight-ready software is provided on the flight computer on the LAICE CubeSat to upload the correct uplink sequence to the LIIB. / Master of Science

Velocity and pressure fluctuations were measured in a wind tunnel in the plane of symmetry in front of a wing-body junction at an approach free-stream velocity U_ref=28.3 m/s and momentum thickness Reynolds number Re_θ=6900. The cylindrical wing shape was a 3:2 elliptic nose attached to a NACA 0020 tail at maximum thickness. A two component laser Doppler anemometer was used to make the velocity measurements. Pressure measurements were made at two locations upstream of the wing by a pair of Sennheiser microphones. A relationship between the velocity and pressure was determined to reveal the bimodal structure of the flow field. Histograms and power spectra of both quantities are presented along with cross-spectra and cross-correlations. The velocity power spectra revealed spectral slopes of -1 through out the entire flow field. Velocity and pressure autospectra show the bimodal region to be dominated by low frequency fluctuations centered at ft/U_ref=5x10⁻², where f is the average frequency and t is the maximum wing thickness. Coherence was found between velocity fluctuations and pressure fluctuations in three frequency bands, each associated with a particular region of the wing-body junction flow field. Low frequency coherence between the velocity and pressure was found in the vicinity of the junction vortex where large scale unsteady bimodal structures are formed. High frequency coherence dominates closer to the juncture of the wing and body. Conditionally-averaged velocity vectors were used to help identify the flow structure in the juncture. These vectors were used along with the other data to propose a model consisting of a single vortex rolling up and moving downstream. As the vortex moves downstream, it wraps around the wing which causes it to stretch and eventually dissipate. Preceding the vortex roll up, an intake of fluid down the wing occurs which is believed to be the cause of the bimodal unsteadiness. / Master of Science

Traditionally, Retarding Potential Analyzers (RPAs) operate exclusively on large satellites due to the size, power, and mass constraints posed by nano-satellites like CubeSats. These sensors take in-situ measurements of Earth's atmospheric ion current during a range of time-varied ``retarding" voltage steps. Curve-fitting the retarding voltage versus collected current data provides derived measurements of ion density, ram velocity, and temperature. In order to successfully miniaturize these instruments and validate their performance prior to launch, thorough calibration and comprehensive end-to-end testing must be performed. This paper discusses the difficulties of performing complete system validation in ground-based vacuum chamber testing for RPAs. A procedure for RPA instrument calibration will be presented along with the calibration results for the Lower Atmosphere/Ionosphere Coupling Experiment (LAICE) CubeSat RPA. This paper presents a user-friendly and robust software control suite developed to read, parse, and interpret the data from the LAICE RPA. Electronics noise testing and analysis defines the performance boundaries of the instrument electronics. End-to-end testing of the LAICE RPA with a hot-filament ion source simulating the space plasma verifies the function of the LAICE RPA sensor and electronics, as well as the software control, thus qualifying the instrument for on-orbit use. / Master of Science

Active magnetic bearings have several potential applications in spacecraft design. Based on the gyrostat model, we develop equations that describe the internal torques and forces that occur between the body and one of the attached wheels. We evaluate the transverse torques for the torque--free gyrostat and a gyrostat undergoing attitude maneuvers using momentum wheels and external torques. We then apply these internal forces to a model of an active magnetic bearing system and discuss their effects on the force limit, the actuator slew rate and the equivalent stiffness and damping parameters. As a basis for this study we use the Distributed Spacecraft Attitude Control System Simulator (DSACSS) with a Revolve MBRotor active magnetic bearing system. The results of several numerical simulations show that the magnitude and frequency of the internal torques remain small over the estimated range of motion of the DSACSS--MBRotor gyrostat. As such, the transverse torques caused by the rotational motion remain less than the discussed performance limits. We show that the magnitude of the internal torques can also be minimized by reducing the axial moment of inertia of the wheel. Furthermore, we discuss the equivalent Jeffcott model. By applying a standard Proportional--Integral--Derivative controller to the active magnetic bearing both the equivalent stiffness and damping parameters remain constant. / Master of Science

The purpose of this thesis is to develop an existing design of a fully controllable multi-rotor vehicle toward simulating small satellite dynamics, enabling technology development to be accelerated and component failure risks to be mitigated by providing a testing platform with dynamics similar to those of small satellites in orbit. Evaluating dynamics-sensitive software and hardware components for use in small satellite operations has typically been relegated to simulated or physically constrained testing environments. More recently, researchers have begun using multi-rotor aerial vehicles to mimic the orbital motion of such satellites, further increasing simulation fidelity. The dynamical nature of multi-rotor vehicles allows them to accurately simulate the translational dynamics of a small satellite, but they struggle to accurately simulate rotational dynamics, as conventional multi-rotor vehicles' translational and rotational dynamics are coupled. In this thesis, an optimal design for a multi-rotor vehicle independently controllable in all six degrees of freedom is evaluated as a suitable simulation platform. The design of the proposed physical system is discussed and progress toward its construction is demonstrated. To facilitate future research endeavors, a simulation of the vehicle in a software-in-the-loop environment, using the Gazebo dynamics simulator, is developed and its performance evaluated. This simulation is then used to evaluate the vehicle's feasibility as a small-satellite dynamics simulator by tasking it with tracking dynamic position and attitude time histories representative of a small satellite. / Master of Science / When developing a spacecraft, it can be difficult to accurately test software and hardware that are sensitive to the spacecraft's motion. This difficulty arises because the space environment experienced by orbiting spacecraft allows them to move and rotate freely, and recreating this freedom of motion on earth requires large, expensive, and difficult-to-access test equipment. To make this testing more accessible, researchers have begun using quadcopter drones to mimic some aspects of a spacecraft's motion. While quadcopters can move like an orbiting spacecraft can, their designs do not allow them to rotate like an orbiting spacecraft can, thus providing an incomplete recreation of spacecraft motion. To correct this shortcoming, an existing drone design that is able to move and rotate simultaneously without fear of crashing is developed, with progress shown toward its construction. A software simulation of the drone is developed to help future researchers test software and algorithms before flying it on the physical drone. The simulation is then used to see how well the drone design can recreate the motions that a small spacecraft would experience.

A novel torsion pendulum thrust test stand for micro-Newton-scale spacecraft thrusters is described. The stand is designed to be robust against electromagnetic interference effects internal or external to the thruster being tested. The design and testing of a gas-dynamic calibration thruster is included. This thruster is fully self-contained on the pendulum arm, with no external wires or feedlines connected to the device and impacting the dynamic response of the underlying pendulum. Initial calibration results are shown. Zero drift and hysteresis are present in the results, evidenced by a constant steady-state displacement drift and a return to a different displacement after shutdown of the calibration thruster. Results are compared to theoretical solutions of the equation of motion. An external forcing function of facility effects is described for discrepancies between results and the theoretical solution. Further work to eliminate these effects and add damping are proposed. / MS / Many recently proposed space missions require very fine vehicle attitude and position control in support of their science objectives. Thrusters with the ability to provide this control are currently in development, from laboratory proofs of concept to initial test flights on pathfinding missions. The low levels of thrust produced by these devices, in the range of less than the weight of a mosquito, require specialized test stands with very fine resolution. This thesis describes a novel torsion pendulum design for measuring these thrusters as well as initial validation results from its calibration system using rarefied gas flow. This calibration device is fully-contained on the device’s arm, removing many common sources of compensation factors which are often needed for other test stand designs. A custom-built displacement measuring system for determining angular motion of the pendulum arm is described which allows for measuring angular displacements of the arm to the level of arcseconds and potentially fractions thereof. Initial results suggest measurement of the expected levels of thrust, while some work remains to remove lingering sources of error and achieve more precise thrust data.

Attitude control laws that use control moment gyros (CMGs) and momentum wheels are derived with nonlinear techniques. The control laws command the CMGs to provide rapid angular acceleration and the momentum wheels to reject tracking and initial condition errors. Numerical simulations of derived control laws are compared. A trend analysis is performed to examine the benefits of the derived controllers. We describe the design of a CMG built using commercial off-the-shelf (COTS) equipment. A mixed attitude control strategy is implemented on the spacecraft simulator at Virginia Tech. / Master of Science

The problem of magnetic control for three-axis stability of a spacecraft is examined. Two controllers, a proportional-derivative controller and a constant coefficient linear quadratic regulator, are applied to the system of equations describing the motion of the spacecraft. The stability of each is checked for different spacecraft configurations through simulations, and the results for gravity-gradient stable and non gravity-gradient stable spacecraft are compared. An optimization technique is implemented in an attempt to obtain the best performance from the controller. For every spacecraft configuration, a set of gains can be chosen for implementation in the controller that stabilizes the linear and nonlinear equations of motion for the spacecraft. / Master of Science