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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https://finding-aids.library.gatech.edu/agents/corporate_entities/106

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Publication Search Results

Now showing 1 - 10 of 40

Conceptual thermal response modeling, testing, and design of flexible heatshield insulation materials

(Georgia Institute of Technology, 2018-01-12) Rossman, Grant Andrew

Flexible Thermal Protection Systems (FTPS) have been investigated to support many applications, including thermal protection of inflatable atmospheric entry vehicles. This flexible blanket is composed of a stack of material sheets, including heat rate resistant outer fabrics, heat load resistant insulation, and an air-tight gas barrier to prevent pressure leaks. This dissertation advances the state-of-the-art of thermal modeling, material property testing, and design of FTPS. In this investigation, a one-dimensional (1D) thermal response model is used to predict in-depth temperatures of FTPS layups during arc-jet ground testing. An extended inverse multi-parameter estimation methodology is developed to improve thermal model prediction accuracy. This method utilizes concepts from inverse heat transfer analysis, parameter estimation, and probabilistic analysis. Thermal response model input parameters are adjusted to minimize the error between temperature predictions and in-depth temperature measurements from arc-jet ground testing. Some FTPS insulators experience decomposition under extreme heating conditions, while others do not. In this investigation, a thermogravimetric analysis (TGA) experimental campaign was designed and executed to further characterize fibrous insulators that undergo decomposition. This material testing methodology was developed to obtain the approximate distribution of activation energy. Associated activation energies were inserted into corresponding thermal response models to improve temperature prediction accuracy. In this investigation, a simulation-based FTPS insulator design methodology is developed to obtain a final FTPS insulator configuration. This design process uses inputs such as candidate insulators, insulator material properties, and a nominal mission profile. Candidate insulators are designed efficiently using an improved thermal response model, providing FTPS insulator stackup configurations that satisfy mission requirements.
Use of the Mars atmosphere to improve the performance of supersonic retropropulsion

(Georgia Institute of Technology, 2017-05-23) Gonyea, Keir C.

NASA has landed seven vehicles on the surface of Mars using parachutes for supersonic descent. These parachutes are unsuited to future high mass missions due to inflation, drag, and aerothermodynamic complications. Supersonic retropropulsion is a candidate technology to replace supersonic parachutes, but is hindered by its large associated propellant mass. Atmospheric-breathing propulsion systems may reduce this mass constraint by ingesting oxidizer from the surrounding atmosphere. However, the Martian atmosphere, which is composed of primarily carbon dioxide, necessitates that metal fuels be used in order to combust the available oxidizer. This thesis advances the state of the art of atmospheric-breathing supersonic retropropulsion (ABSRP) by providing the first exploration into the feasibility and potential performance of ABSRP as a technology solution for high-mass Mars missions. Specific advancements include the development of modeling methods and tools, the evaluation of conceptual ABSRP performance and sensitivities, and the formulation of vehicle concepts. Model development targeted components and subsystems most relevant to ABSRP in order to capture the necessary physics and provide a preliminary integrated vehicle simulation for future conceptual design efforts. Models were developed to assess metal – CO2 combustion performance and sensitivity to both the engine design and operating regime. These tools include an equilibrium combustion simulation to evaluate engine efficiency, a finite-rate kinetics simulation to investigate the time-dependent phenomena, and a particle burning simulation to assess diffusion effects. Case studies are presented for ABSRP relevant mixtures and conditions to predict propulsion performance of the ABSRP engine across a range of conditions and verify that reasonably sized combustion chambers can provide complete combustion of the propellant. Exploration of the performance results indicate that ABSRP systems have promising propulsive performance relative to comparative rocket systems and do not have unacceptable burning timescale constraints. The propulsion system results are used in an ABSRP vehicle model, which accounts for the variable engine performance across different flight regimes. This model is used to search the design space and determine the performance and sensitivity of multiple proposed ABSRP vehicle concepts relative to competing propulsive solutions. The investigation includes an assessment of feasible and unfeasible regions of the design space in addition to design trends for optimal configurations. Mass favorable vehicles of multiple architectures are compared to understand their relative performance. Vehicle architectures involving ABSRP are seen to have optimal mass performance, which demonstrates the potential applicability of atmospheric-breathing propulsion for Mars descent.
Development of an Earth SmallSat Flight Test to Demonstrate Viability of Mars Aerocapture

(Georgia Institute of Technology, 2017-05-01) Werner, Michael S.

A smallsat mission concept is developed to demonstrate the feasibility of an aerocapture system at Earth. The proposed mission utilizes aerocapture to transfer from a GTO rideshare trajectory to a LEO. Single-event drag modulation is used as a simple means of achieving the control required during the maneuver. Low- and high-fidelity guidance algorithm choices are considered. Numeric trajectory simulations and Monte Carlo uncertainty analyses are performed to show the robustness of the system to day-of-flight environments and uncertainties. Similar investigations are performed at Mars to show the relevance of the proposed mission concept to potential future applications. The spacecraft design consists of a 24.9 kg vehicle with an attached rigid drag skirt, and features commercially-available hardware to enable flight system construction at a university scale. Results indicate that the proposed design is capable of targeting the desired final orbit, surviving the aerothermodynamic and deceleration environments produced during aerocapture, and downlinking relevant data following the maneuver
Trajectory Trade-space Design for Robotic Entry at Titan

(Georgia Institute of Technology, 2017-05-01) Roelke, Evan

In recent years, scientific focus has emphasized other ocean worlds such as Europa, Enceladus, and Titan, due to their potential for harboring life. The only spacecraft ever to land on these moons was the Huygens Probe in 2005; however, this probe’s main purpose was to study the atmosphere and surface of Titan, with no real landing target. Future missions to other ocean worlds would likely require a science target and thus add several constraints to the mission such as arrival time, entry state, and aeroshell geometry, among others. Of the three ocean worlds previously mentioned, Titan is an optimal target for initial mission concepts for several reasons. The atmospheric composition, winds, and surface features are well studied by Cassini and the Huygens Probe. Additionally, of the aforementioned moons, Titan does not have a thick ice sheet to penetrate in order to sample the surface and/or liquid seas, enabling such mission to double as a stepping stone for missions to other ocean worlds. Finally, Titan exhibits a myriad of interesting planetary features that, if studied, could further the understanding of both Titan’s and the solar system’s geologic history. In this paper we analyze the trade-spaces of various important parameters involved in Entry, Descent, and Landing (EDL) as it pertains to robotic missions for Titan in order to provide a guideline for optimizing a mission’s system parameters while minimizing both system complexity and the landing footprint. It is found that the ideal geometry is a ballistic spherecone body entering from orbit to allow flexibility in the entry state vector. The aerothermodynamic environment is most affected by the entry velocity and the vehicle bluntness ratio, while the peak deceleration is most influenced by the entry velocity and entry flight path angle. In addition, multiple parachutes decrease the landing footprint, impact speed, and descent time compared to single parachute systems, at the expense of being more complex. Larger ballistic coefficients decrease the landing footprint and descent time while increasing the impact speed. Finally, it is discovered that the uncertainty in the entry altitude and flight path angle have the most impact on the final state vector.
Mechanical Design of a Cubesat Aeroshell for an Earth Demonstration of Single-Stage Drag Modulated Aerocapture

(Georgia Institute of Technology, 2016-08-01) Woollard, Bryce A.

The following article documents the conceptual study of a smallsat entry vehicle to be implemented for demonstration of single-stage drag modulated aerocapture at Earth. The specific nature of the contents below focuses on the mechanical design and analysis of the aeroshell and drag device, as well as the mechanisms by which all parts are to be manufactured, assembled and actuated in order to perform the intended orbital maneuver. The results of this study show that accomplishing aerocapture with a cubesat entry vehicle appears to be feasible with a 2U payload and would require approximately 20 kg and 0.1 m3 of secondary payload mass and volume, respectively. First order stagnation point thermal protection sizing suggests that 4.2 cm of PICA would be required globally around the vehicle, although potential exists to optimize this value relative to geometric location. Static stability analysis indicates that the designed vehicle is nose-forward stable for a majority of the atmospheric interface with outstanding questions pertaining to atmospheric egress. Manufacturing costs for a full scale aeroshell would be approximately $15,000 and require roughly 2 months of lead time, dependent on presently available machine shop capabilities.
Mechanical property determination for flexible material systems

(Georgia Institute of Technology, 2016-04-12) Hill, Jeremy Lee

Inflatable Aerodynamic Decelerators (IADs) are a candidate technology NASA began investigating in the late 1960’s. Compared to supersonic parachutes, IADs represent a decelerator option capable of operating at higher Mach numbers and dynamic pressures. IADs have seen a resurgence in interest from the Entry, Descent, and Landing (EDL) community in recent years. The NASA Space Technology Roadmap (STR) highlights EDL systems, as well as, Materials, Structures, Mechanical Systems, and Manufacturing (MSMM) as key Technology Areas for development in the future; recognizing deployable decelerators, flexible material systems, and computational design of materials as essential disciplines for development. This investigation develops a multi-scale flexible material modeling approach that enables efficient high-fidelity IAD design and a critical understanding of the new materials required for robust and cost effective qualification methods. The approach combines understanding of the fabric architecture, analytical modeling, numerical simulations, and experimental data. This work identifies an efficient method that is as simple and as fast as possible for determining IAD material characteristics while not utilizing complicated or expensive research equipment. This investigation also recontextualizes an existing mesomechanical model through validation for structures pertaining to the analysis of IADs. In addition, corroboration and elaboration of this model is carried out by evaluating the effects of varying input parameters. Finally, the present investigation presents a novel method for numerically determining mechanical properties. A sub-scale section that captures the periodic pattern in the material (unit cell) is built. With the unit cell, various numerical tests are performed. The effective nonlinear mechanical stiffness matrix is obtained as a function of elemental strains through correlating the unit cell force-displacement results with a four node membrane element of the same size. Numerically determined properties are validated for relevant structures. Optical microscopy is used to capture the undeformed geometry of the individual yarns.
Supersonic Vehicle Configuration Transitions to Enable Supersonic Retropropulsion during Mars Entry, Descent, and Landing

(Georgia Institute of Technology, 2016-02-29) Blette, David J.

This paper investigates types of supersonic vehicle configuration transition events nec essary to initiation supersonic retropropulsion as part of human-class Mars entry, descent, and landing. This research assumes an entry vehicle with a 105 mT entry mass and an ellipsled aeroshell similar to the NASA EDL Design Reference Architecture 5.0. All entry architectures are assumed all-propulsive. Three transition architectures are considered: a pitch-around maneuver, an aeroshell front-exit, and an aeroshell hinged-exit. Propulsive subsystem thrust requirements are defined for the pitch-around maneuver. For transitions involving solid mass ejections, debris flight envelopes are determined and compared to a descent vehicle trajectory under a modified gravity turn. It is shown that far-field recon tact risks exist for the proposed architectures involving solid mass ejections and recontact mitigation schemes are required.
Entry Characteristics of a Half-Ogive Aeroshell at Earth

(Georgia Institute of Technology, 2016-02-05) Booher, Robert M.
Automated trajectory control for proximity operations using relative orbital elements

(Georgia Institute of Technology, 2015-04-02) Spencer, David Allen

This dissertation develops a methodology for automated trajectory control of a spacecraft about a non-maneuvering target. The methodology utilizes relative orbital elements (ROEs), combined with guidance laws based upon artificial potential functions (APFs), to perform automated trajectory planning and maneuver design. The investigation provides a definitive reference on the definition and use of ROEs for relative proximity operations. The detailed derivation of ROEs is provided, along with transformations between ROEs and relative Cartesian state elements, characteristics of unforced motion in terms of ROEs, and the effect of impulsive maneuvers on ROEs. Operationally-useful guidance algorithms utilizing ROEs are developed and demonstrated. These ROE-based algorithms for rendezvous, circumnavigation and station-keeping provide a toolkit for relative proximity operations mission planning. A new approach for APF formulation using ROEs as the target variables is developed. While previous approaches allowed targeting of a specified relative position, the present approach allows the targeting of relative orbit geometries. The approach capitalizes upon the orbital dynamics represented through the ROEs, and retains the computational simplicity offered by the APFs. Formulations for the APF targeting of individual ROEs, as well as simultaneous targeting of a set of ROEs, are established. An approach for combining ROE targeting using APFs with obstacle avoidance is presented. The trajectory guidance algorithm performance is evaluated using a flight-like guidance, navigation and control simulation environment, including orbital perturbations. Algorithm performance is established through a set of operationally relevant scenarios. The guidance algorithms are shown to be capable of correcting for environmental disturbances, while meeting the targeted relative orbits in an automated fashion.
Improved analytical methods for assessment of hypersonic drag-modulation trajectory control

(Georgia Institute of Technology, 2015-04-01) Putnam, Zachary Reed

During planetary entry, a vehicle uses drag generated from flight through the planetary atmosphere to decelerate from hyperbolic or orbital velocity. To date, all guided entry systems have utilized lift-modulation trajectory control. Deployable aerodynamic devices enable drag-modulation trajectory control, where a vehicle controls its energy and range during entry by varying drag area. Implementation of conventional lift-modulation systems is challenging for deployable systems. In contrast, drag-modulation trajectory control may be simpler and lower-cost than current state-of-the-art lift-modulation systems. In this investigation, a survey of analytical methods for computing planetary entry trajectories is presented and the approximate analytical solution to the entry equations of motion originally developed by Allen and Eggers is extended to enable flight performance evaluation of drag-modulation trajectory control systems. Results indicate that significant range control authority is available for vehicles with modestly sized decelerators. The extended Allen-Eggers solution is closed-form and enables rapid evaluation of nonlifting entry trajectories. The solution is utilized to develop analytical relationships for discrete-event drag-modulation systems. These relationships have direct application to onboard guidance and targeting systems. Numerical techniques were used to evaluate drag-modulation trajectory control for precision landing and planetary aerocapture missions, including development of prototype real-time guidance and targeting algorithms. Results show that simple, discrete-event drag-modulation trajectory control systems can provide landed accuracies competitive with the current state of the art and a more benign aerothermal environment during entry for robotic-scale exploration missions. For aerocapture, drag-modulation trajectory control is shown to be feasible for missions to Mars and Titan and the required delta-V for periapsis raise is insensitive to the particular method of drag modulation. Overall, results indicate that drag-modulation trajectory control is feasible for a subset of planetary entry and aerocapture missions. To facilitate intelligent system selection, a method is proposed for comparing lift and drag-modulation trajectory control schemes. This method applies nonlinear variational techniques to closed-form analytical solutions of the equations of motion, generating closed-form expressions for variations of arbitrary order. This comparative method is quantitative, performance-based, addresses robustness, and applicable early in the design process. This method is applied to steep planetary entry trajectories and shows that, in general, lift and drag-modulation systems exhibit similar responses to perturbations in environmental and initial state perturbations. However, significant differences are present for aerodynamic perturbations and results demonstrate that drag systems may be more robust to uncertainty in aerodynamic parameters. Finally, the results of these contributions are combined to build a set of guidelines for selecting lift or drag-modulation for a Mars Science Laboratory-class planetary entry mission.