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Space Systems Design Laboratory (SSDL)

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https://hdl.handle.net/1853/70747

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Daniel Guggenheim School of Aerospace Engineering

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https://finding-aids.library.gatech.edu/agents/corporate_entities/1138

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Publication Search Results

Now showing 1 - 10 of 12

Trade Study and Analysis for a Deployable Drag Device for Launch Vehicle Upper Stage Deorbit

(Georgia Institute of Technology, 2015-05-02) Long, Alexandra C.

Orbital debris is a growing problem in low Earth orbit; it has crossed a threshold of critical density where the number of debris objects will grow exponentially unless mitigated. Spent launch vehicle upper stages represent a problematic category of orbital debris in highly utilized orbits. They can stay in orbit for well over 100 years if left to deorbit naturally, and they represent a significant fraction of large space debris in low-Earth orbit. It is estimated that removing a few large objects per year will mitigate the exponential growth of debris. To address the debris problem, a trade study was conducted to determine a deployable drag device to accelerate the orbit degradation of upper stages. Following the operation of the upper stage, the drag device will be deployed to decrease the orbit lifetime of the system. The design is targeted toward upper stages launched into orbital altitudes ranging from 650-850 km. Three categories of deployable drag devices are being investigated: drag sails, inflatable aerodynamic decelerators, and electrodynamic tethers. These are compared to the option of using residual propellant in the upper stage to perform a burn to initiate a deorbit trajectory. The device will be mounted to the upper stage using a standardized secondary payload launch interface, such as a CubeSat deployer device or the EELV Secondary Payload Adapter (ESPA). The trade study compared the drag device configurations based on cost, risk, and deorbit time. A maximum deorbit period of 25 years is a performance design requirement. The propulsive option was shown to be the lowest cost option, however the drag device is more mass efficient and has less of an impact to the payload capability of the launch vehicle. A drag sail design is proposed as a baseline design for the device. A stability analysis was conducted to determine the configuration of the device, then the initial baseline design with preliminary component selection and an initial structural analysis were investigated
Georgia Tech Small Satellite Real-Time Hardware-in the-Loop Simulation Environment: SoftSim6D

(Georgia Institute of Technology, 2015) Chait, Sean B.

The capabilities of small satellites produced by the university and small business community have seen a sharp rise in recent years. With this growth in capabilities has come an increase in mission complexity to encompass those architectures previously only found in well-funded government programs, including proximity operations. The inherent complexity of proximity operations-based missions introduces a great deal of risk to the mission’s success. The low-budget nature of the small satellite community has limited the development of relevant testing infrastructure to match the pace of mission complexity increase to adequately mitigate risk. This research will leverage the standardization of CubeSat components to develop a highly adaptable hardware-in-the-loop testing capability for the verification and validation of small satellite avionics boards and flight software. MATLAB© Simulink Real-Time will be utilized to create a user friendly framework that can easily be adapted to support a wide range of small satellite mission architectures. This architecture, known as SoftSim6D, has been designed to thoroughly exercise the robustness of a satellite with the primary aim of minimizing mission risk to ensure full mission success. An examination of the overall framework, verified capabilities, and current variants will be discussed.
Prox-1 Guidance, Navigation, & Control Overview: Development, Algorithms, and Integrated Simulation

(Georgia Institute of Technology, 2014-12-01) Schulte, Peter Z.

This report describes the development and validation process of a highly automated Guidance, Navigation, & Control (GN&C) subsystem for a small satellite on-orbit inspection application. The resulting GN&C subsystem performs proximity operations (ProxOps) without human-in the-loop interaction. The report focuses on the description of the GN&C algorithms, the integration and testing of GN&C software, and the development of decision logic to address the question of how such a system can be effectively implemented for full automation. This process is unique because a multitude of operational scenarios must be considered and a set of complex interactions between various GN&C components must be defined to achieve the automation goal. The GN&C subsystem for the Prox-1 satellite is currently under development within the Space Systems Design Laboratory at the Georgia Institute of Technology. The Prox-1 mission involves deploying the LightSail 3U CubeSat, entering into a leading or trailing orbit of LightSail using ground-in-the-loop commands, and then performing automated ProxOps through formation flight and natural motion circumnavigation maneuvers. Operations such as these may be utilized for many scenarios including on-orbit inspection, refueling, repair, construction, reconnaissance, docking, and debris mitigation activities. Prox-1 uses onboard sensors and imaging instruments to perform its GN&C operations during on-orbit inspection of LightSail. Navigation filters perform relative orbit determination based on images of the target spacecraft, and guidance algorithms conduct automated maneuver planning. A slew and tracking controller sends attitude actuation commands to a set of control moment gyroscopes, and other controllers manage desaturation, detumble, and target acquisition/recovery. All Prox-1 GN&C components are developed in a MATLAB/Simulink six degree-of-freedom simulation environment and are integrated using decision logic to autonomously determine when certain actions should be performed. The complexity of this decision logic is the main challenge of this process, and the Stateflow tool in Simulink is used to establish logical relationships and manage data flow between each of the individual GN&C hardware and software components. Once the integrated GN&C simulation is fully developed in MATLAB/Simulink, the algorithms are autocoded to C/C++ and integrated into flight software. The subsystem is tested using hardware-in-the-loop on the flight computers and other hardware
Structural Design, Analysis, and Test of the Prox-1 Spacecraft

(Georgia Institute of Technology, 2012-12-01) Willingham, Allison L.

HE Prox-1 spacecraft is Georgia Institute of Technology’s entry into the 7th University Nanosatellite Program Competition, a two year cycle competition for the AFRL where university teams consisting of both graduate and undergraduate students design, build, and test a 50 kg nanosatellite for a team-specified mission. Judging is based on various presentations to the AFRL review teams, importance of the mission to AFRL objectives, and development of a sound nanosatellite system among other criteria [5]. Prox-1 is a nanosatellite which will demonstrate the use of low-thrust propulsion for automated safe trajectory control during proximity operations. Passive, image based observations will be used for the navigation and closed loop attitude control of Prox-1 relative to a deployed CubeSat. Prox-1’s objectives include: Rendezvous and proximity operations with a target CubeSat, automated relative navigation and trajectory control, closed-loop attitude control based upon automated image processing, and relative orbit determination using image-based angle and range estimates, validated by the Mission Operations System [4]. The student’s particular research involved design, build, and test of the structural components of the Prox-1 satellite. This paper will describe what design information was based on previous Prox-1 structure iterations, what design modifications were made to improve the structure’s capabilities and meet requirements, what analysis and testing was performed to validate those requirements, and what was needed to integrate with the subsystem components. When referring to different plate orientations in this document, the Prox-1 body coordinate frame is used. This is centered at the middle of the Lightband interface ring on the bottom plate, and in the same plane as the Launch Vehicle Interface. In the final structure configuration, the X-axis is pointing toward the Ppod deployment direction and cameras, the positive Y-axis is in the direction opposite of the thruster, and the Z-axis is pointing from the LVI plate toward the top plate [2]. All figures depicting the spacecraft will have this body coordinate frame pictured
Design and Analysis of the Deorbit and Earth Entry Trajectories for SPORE

(Georgia Institute of Technology, 2012-05-02) Nehrenz, Matthew

Small Probes for Orbital Return of Experiments (SPORE) provides on-orbit operation and recovery of small payloads. The flight system architecture consists of a service module for on-orbit operations and deorbit maneuvering, and an entry vehicle for atmospheric entry, descent, and landing. Prior to approximating a landing footprint with a Monte Carlo analysis on the entry trajectory, the entry state uncertainties must be characterized. These uncertainties arise from errors induced by the guidance system and thruster pointing control during the deorbit maneuver. In order to capture the effect that these errors have on the entry state uncertainty, the service module’s attitude determination and control system (ADCS) and guidance system were both modeled in Matlab. By incorporating the ADCS loop into the guidance loop, the effect of pointing errors during the deorbit trajectory combined with errors in the guidance system can be assessed. A Monte Carlo analysis is performed on this 3+3 DOF deorbit simulation (which terminates at entry interface), resulting in an entry state covariance. The analysis is performed on the three orbits under consideration for SPORE: ISS, LEO, and GTO. Finally, the resulting entry state covariance from the deorbit simulation is used as input for an entry, descent, and landing trajectory Monte Carlo analysis. Landing footprint, heating, and g-loading are considered for trajectories targeting Woomera Test Range in Australia
Design, Analysis, and Testing of the Prox-1 Satellite Structure

(Georgia Institute of Technology, 2011-12-01) Pappas, Thomas D.
Prox-1 Attitude Determination and Control

(Georgia Institute of Technology, 2011-12-01) Pietruszewski, Amanda N.

The Prox-1 mission is Georgia Institute of Technology’s entry to the University Nanosat Program’s competition cycle. Since the goal of the program is to design a complete satellite that will function on orbit, it must have fully designed and space capable subsystems. This paper details the design of the Attitude Determination and Control Subsystem. First, the requirements for the subsystem are derived from the mission requirements, and then trades are presented to find which architecture is the best selection to meet these requirements. Then the trades for specific hardware selections are done, followed by the hardware budgets and tests. Once the hardware is chosen, the algorithms can be explored and described. A rough analysis of detumble length is done, followed be a detailed derivation and explanation of the coarse attitude control system. A filter for attitude determination is explained, followed by a discussion of the fine attitude determination and control system and then the flight rules for the system. Thus, the design of the system done to date is presented in this paper
Near-earth object extravehicular activities: Using apollo and iss operations to map low-gravity terrestrial spacewalk objectives and challenges

(Georgia Institute of Technology, 2011-12-01) Gast, Matthew A.

The notion of human exploration of a near-Earth object (NEO) is nothing new. Jules Verne wrote about this very idea in his story “Off on a Comet,” first published in France in 1877. Since that time, a number of studies have examined NEO exploration for scientific purposes, in-situ resource utilization, mineralogical exploitation and even planetary defense; as early as 1966, a study was conducted to utilize the Apollo program hardware to fly by asteroid Eros 433 [32]. Yet there is very little in the literature archive addressing extra-vehicular activities operations on the surface of a near-Earth object. The arguments for manned missions to near-Earth objects have been presented in a number of papers, recognizing astronauts’ adaptability to real-time challenges, the capability to collect geological samples while identifying the overall geological context, and the ability to return a great quantity of those geological samples to Earth, as just a few of the many reasons for a NEO manned mission. Few studies, however, have identified or discussed the myriad challenges of performing surface operations in an environment where the gravitation is considerably less than that of the Moon, but not negligible like the micro-gravity of an International Space Station (ISS) – based EVA. Using the operational experience learned from NASA’s various human exploration programs, this paper will identify key challenges unique to NEO surface operations. Furthermore, this paper will map the applicable EVA tasks from both the Apollo program’s lunar exploration missions and ISS construction to present an EVA operational concept for NEO surface exploration. Through mapping the applicable Apollo and ISS tasks to the surface of a NEO, relevant operational objectives and challenges are identified, and conceptual approaches to meeting the NEO EVA mission objectives and mitigating key risks are discussed
Center For Space Systems Mission Success Assurance: Requirement Verification and Spacecraft Integration and Test Program

(Georgia Institute of Technology, 2011-05-01) McNeese, Sarah E.

This document provides the framework of mission success assurance for space-flight projects developed within the Center for Space Systems (CSS). These guidelines, protocols, and procedures were defined during the development of the first project within the laboratory: the Rapid Reconnaissance and Response (R3 ) small satellite mission. The document first details the guidelines and systems put in place for requirement development and verification, the integration and test program, and the hardware development best practices. It then enters into a discussion of how this framework was implemented for R3 , including specific details and examples of the tests planned and executed. The Requirement Verification Matrix (RVM) developed for R3 is provided as an Appendix in both electronic and hard copy as a model. Templates of all documentation developed to support CSS flight projects are also provided as an Appendix in both electronic and digital versions. All documentation of the R3 verification and testing program (hardware inspections, test planning forms, test completion records, etc.) is provided in an electronic appendix that is provided alongside this document.
Thermal Algorithms

(Georgia Institute of Technology, 2010-05-01) Vedie, Nathalie

An overview of the thermal algorithms utilized in the R3 mission is presented. The R3 satellite will have thermal algorithms to process images taken by the thermal imager. The algorithms will calculate the area of features based on whether the feature’s size matches the intended features area and on whether the temperature of the feature falls within a certain range. The center of each matching feature can be calculated and validated onboard the satellite. The algorithms will also be able to detect edges to various ends.