Organizational Unit:

Daniel Guggenheim School of Aerospace Engineering

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https://hdl.handle.net/1853/70742

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College of Engineering

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Aerospace Design Group

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Aerospace Systems Design Laboratory (ASDL)

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Air Transportation Laboratory

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Cognitive Engineering Center

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Space Systems Design Laboratory (SSDL)

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Publication Search Results

Now showing 1 - 10 of 96

Deployable Drag Device for Launch Vehicle Upper Stage De-orbit

(Georgia Institute of Technology, 2014-09) Long, Alexandra C. ; Spencer, David A.

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. An aerostable drag sail design is proposed as a baseline design for the device.
Development Of An Integrated Spacecraft Guidance, Navigation, and Control Subsystem For Automated Proximity Operations

(Georgia Institute of Technology, 2014-09) Shulte, Peter Z. ; Spencer, David A.

This paper 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 paper focuses on 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, thruster firing, 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.
Cooperative Scenarios for Human Exploration Beyond Low Earth Orbit

(Georgia Institute of Technology, 2014-09) Battat, Jonathan ; Alifanov, Oleg ; Braun, Robert D. ; Crawley, Edward ; Logsdon, John ; Zeleny, Lev ; Borowitz, Mariel ; Capparelli, Emanuele ; Davison, Peter ; Golkar, Alessandro ; Steinfeldt, Bradley A.

There is an international need to define a concrete strategy and plan to implement that strategy for the initial human exploration missions beyond Low Earth Orbit (LEO). Across all stakeholders, there is a growing consensus that the long term objective of global human space exploration is the long duration presence of people on the Martian surface. Along the pathway between current activities in LEO and eventual Mars outposts are a variety of preparatory exploration missions and intermediate goals. Over the last decade several different initial steps along these pathways beyond LEO have been proposed. It is important to build international consensus on such a plan soon because future missions require near-term investments for new capabilities with no single nation committing resources to achieve all the steps of an ambitious program on its own. The goal of this work is to enumerate and evaluate scenarios for cooperative missions beyond LEO that achieve incremental development of human exploration capabilities. Towards the goal of generating scenarios for cooperative missions beyond LEO, proposed missions and capabilities from a variety of international actors have been assessed. Presented in this paper are results of a survey of proposed missions and a series of interviews with industry experts knowledgeable about both the technical and geopolitical issues in forging a sustainable path towards Mars. There are four realistic proposals for initial human exploration beyond LEO: a cis-Lunar habitat, asteroid redirect, Mars flyby, and a Lunar surface sortie. In the absence of top-down agreements, such as those governing the International Space Station, that specify partnership responsibilities and privileges, ad-hoc exchanges within individual development projects or for specific mission capabilities is most likely to facilitate international cooperation in the coming years. General LEO transportation logistics and habitation functions are shared by many actors and allow for exchange of services and utilization of exploration assets if designed into the critical path. Given the early stage of readiness, it is possible that subsystem-level coordination could be pursued for an advanced habitation element. Other technologies are either niche (robotics) or have national sensitivities (in-space propulsion) that make them less desirable for subsystem-level coordination.
Pose-Tracking Controller for Satellites with Time-Varying Inertia

(Georgia Institute of Technology, 2014-08) Filipe, Nuno ; Holzinger, Marcus J. ; Tsiotras, Panagiotis

Satellite proximity operations have been identified by NASA and the USAF as a crucial technology that could enable a series of new missions in space. Such missions would require a satellite to simultaneously and accurately track time-varying relative position and attitude profiles. Moreover, the mass and moment of inertia of a satellite are also typically time- varying, which makes this problem even more challenging. Based on recent results in dual quaternions, a nonlinear adaptive position and attitude tracking controller for satellites with unknown and time-varying mass and inertia matrix is proposed. Dual quaternions are used to represent jointly the position and attitude of the satellite. The controller is shown to ensure almost global asymptotic stability of the combined translational and rotational position and velocity tracking errors. Moreover, sufficient conditions on the reference motion are provided that guarantee mass and inertia matrix identification. The controller compensates for the gravity force, the gravity-gradient torque, Earth's oblateness, and unknown constant disturbance forces and torques. The proposed controller is especially suited for satellites with relatively high and quick variations of mass and moment of inertia, such as highly maneuverable small satellites equipped with relatively powerful thrusters and control moment gyros.
Incorporating Uncertainty in Admissible Regions for Uncorrelated Detections

(Georgia Institute of Technology, 2014-08) Worthy, Johnny L., III ; Holzinger, Marcus J.

Admissible region methods for initial orbit determination are generally implemented without considering uncertainty in observations or observer state. In this paper a generalization of the admissible region approach is introduced that more accurately accounts for uncertainty in the constraint hypothesis parameters used to generate the admissible region. Considering the uncertainty to have Gaussian distributions, the proposed relationship between provided information uncertainty and admissible region uncertainty results directly in an analytical approximate probability density function. The methodology is extended to account for admissible regions with multiple overlapping constraint hypothesis. The proposed approach is applied to an example optical detection to demonstrate the quality of the approximation and the sensitivity of the resulting distribution to typical errors.
Extension and Enhancement of the Allen-Eggers Analytic Solution for Ballistic Entry Trajectories

(Georgia Institute of Technology, 2014-06) Putnam, Zachary R. ; Braun, Robert D.

The closed-form analytic solution to the equations of motion for ballistic entry developed by Allen and Eggers is extended and enhanced with a method of choosing an appropriate constant flight-path angle, limits based on the equations of motion and acceptable ap- proximation error are proposed that bound the domain of applicability, and closed-form expressions for range and time-dependency. The expression developed for range to go exhibits error that may low enough for onboard drag-modulation guidance and targeting systems. These improvements address key weaknesses in the original approximate solution. Results show that the extended and enhanced Allen-Eggers solution provides good accuracy across a range of ballistic coefficients entries at Earth with initial flight-path angles steeper than -7 deg.
Propulsion System Design for a Martian Atmospheric Breathing Supersonic Retropropulsion Engine

(Georgia Institute of Technology, 2014-06) Gonyea, Keir C. ; Braun, Robert D.

Design and analysis was performed on an atmospheric breathing propulsion system to land large-scale spacecraft on Mars. Initial feasibility of the engine was investigated analytically by employing equilibrium combustion and finite rate kinetics simulations in addition to 1st order propellant mass and inlet sizing. ISP values (based on total propellant usage) were determined to be on the order of 120s-160s for onboard subsystems having a 10- to-1 oxidizer compression ratio. This corresponds to an ISP of 600s-800s based on fuel consumption. While Mg-CO2 mixtures have significant ignition constraints, favorable conditions were found, yielding ignition delay times of less than 1ms, by simultaneously employing designs exploiting both large reentry Mach numbers and modest compression ratios. These combinations allow for combustion to occur within moderately sized combustion chambers. The 1st order sizing calculations confirmed that atmospheric breathing supersonic retropropulsion has the potential for significant mass savings over traditional retropropulsion architectures. Engines sized with an oxidizer-to-fuel ratio of 4 require half the propellant consumption for an equivalent change in velocity. Inlet capture areas of the examined atmospheric breathing propulsion systems were on the order of the corresponding entry vehicle projected area. Therefore, this study envisioned an annular inlet design, which encircled the vehicle forebody. The aforementioned analyses address some of the challenges that need to be solved in order to ultimately obtain a practical atmospheric breathing supersonic retropropulsion system for Mars descent.
Application of Magnetohydrodynamic Energy Generation to Planetary Entry Vehicles

(Georgia Institute of Technology, 2014-06) Ali, Hisham K. ; Braun, Robert D.

Proposed missions such as a Mars sample return mission and a human mission to Mars require landed payload masses in excess of any previous Mars mission. Whether human or robotic, these missions present numerous engineering challenges due to their increased mass and complexity. To overcome these challenges, new technologies must be developed, and existing technologies advanced. Mass reducing technologies are particularly critical in this effort. The proposed work aims to study the suitability of multi-pass entry trajectories for reclaiming of vehicle kinetic energy through magnetohydrodynamic power generation from the high temperature entry plasma. Potential mission and power storage configurations are explored, with results including recommended trajectories, amount of kinetic energy reclaimed, and additional system mass for various energy storage technologies.
Extensibility of a Linear Rapid Robust Design Methodology

(Georgia Institute of Technology, 2014-01) Steinfeldt, Bradley A. ; Braun, Robert D.

The extensibility of a linear rapid robust design methodology is examined. This analysis is approached from a computational cost and accuracy perspective. The sensitivity of the solution's computational cost is examined by analysing effects such as the number of design variables, nonlinearity of the CAs, and nonlinearity of the response in addition to several potential complexity metrics. Relative to traditional robust design methods, the linear rapid robust design methodology scaled better with the size of the problem and had performance that exceeded the traditional techniques examined. The accuracy of applying a method with linear fundamentals to nonlinear problems was examined. It is observed that if the magnitude of nonlinearity is less than 1000 times that of the nominal linear response, the error associated with applying successive linearization will result in errors in the response less than 10% compared to the full nonlinear error.
Asymmetrically Stacked Tori Hypersonic Inflatable Aerodynamic Decelerator Design Study for Mars Entry

(Georgia Institute of Technology, 2014-01) Harper, Brooke P. ; Braun, Robert D.

The Mars missions envisioned in the future require payload mass in excess of the current capable limit for entry vehicle technology. Deployable Hypersonic Inflatable Aerodynamic Decelerators offer one solution to successfully carry out the beginning of an entry architecture as payload mass increases. The majority of the research that has been conducted on these structures only focuses on axisymmetric geometries. In this investigation, aerodynamic performance and stability is examined for three proposed asymmetric families that can generate non-zero lift-to-drag ratios at 0° angle of attack. Advantages of an asymmetric lifting Hypersonic Inflatable Aerodynamic Decelerator include a larger entry corridor, reduced peak heating, larger range, and improved landing accuracy. In particular, there is potential to increase drag performance and reduce ballistic coefficient to mitigate entry, descent, and landing concerns. Blunt, asymmetric Hypersonic Inflatable Aerodynamic Decelerator designs considered are assembled from stacked tori configurations with a base diameter of 20 m and the capability to interface with a 10 m diameter rigid center body. The configurations considered are capable of producing hypersonic lift-to-drag ratios between ~0.6 and ~0.1 for angles of attack ranging from -30° to 20°. A 40 (t) entry mass, approximate mass of large robotic or human scale mission, is assumed resulting in ballistic coefficients from ~78 kg/m2 to ~113 kg/m2. From the analyses conducted thus far, encouraging results project asymmetric Hypersonic Inflatable Aerodynamic Decelerators as conceivable candidates for future large scale Mars missions.