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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 22

Evaluation of Deployable Aerosurface Systems for Mars Entry

(Georgia Institute of Technology, 2012-12-14) Cruz-Ayoroa, Juan G.

One of the challenges presented by the exploration of Mars is the entry, descent and landing (EDL) of payloads to the surface. Current robotic missions to Mars are reaching the limist of existing Viking heritage EDL technologies. A number of EDL technology improvements can be made to extend the capabilities beyond the current landed mass limits, including increasing the entry vehicle hypersonic drag and lift capability. Technologies being currently studied include inflatable aerodynamic decelerators, which are designed to increase vehicle drag. Many of these concepts center on axisymmetric designs, which provide high drag but relatively low lift and are most easily integrated to blunt entry vehicles. However, due to packaging density and launch vehicle fairing constraints, it is likely that future missions will require the use of slender bodies. This study investigates three deployable concepts designed to provide better integration into a slender vehicle while augmenting its performance by increasing its hypersonic drag. The deployable aerosurfaces are applied to a 5 meter diameter slender vehicle for a robotic mission at Mars with entry masses ranging from 10 to 60t. A multidisciplinary design optimization framework is used to estimate the landed mass capability of each system. Results show that the deployable concepts can significantly improve payload mass capability by reducing the terminal propulsion propellant required. Initial feasibility studies show that the concepts are hypersonically statically stable and comply with mechanical and thermal material capabilities
Implementation of a Mesomechanical Material Model for IAD Fabrics within LS-DYNA

(Georgia Institute of Technology, 2012-12-14) Hill, Jeremy

The implementation and evaluation of a high fidelity material model for dry fabrics is the main objective of this paper. Inflatable Aerodynamic Decelerators (IADs) and other air inflated structures quite often utilize woven fabrics due to their lightweight and high loading carrying capabilities. Design optimization of these inflated structures relies on a detailed understanding of the woven fabric mechanics. Woven fabrics are composite orthotropic materials that respond differently under load from traditional solid mechanics. While low fidelity fabric materials usually assume a continuous medium, a higher fidelity model needs to account for the reorientation of yarns and weave geometry. An existing mesomechanical material model within the LS-DYNAÒ commercial non-linear finite element software package is utilized. In this paper, experimental stress-strain data for Kevlar 129 samples are validated against numerical simulations of models with matching geometry and loading conditions.
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
SPORE Parachute Design and Selection

(Georgia Institute of Technology, 2012-12-01) Stout, Stephanie E.

Eleven parachute types are investigated to determine the best option for the SPORE 1U LEO and 2x2U LEO configurations. The 1U LEO configuration must meet a 5 meters per second impact velocity requirement, and the 2x2U LEO configuration must meet a 40g acceleration limit throughout its trajectory, constraining both the impact deceleration and the parachute opening deceleration due to inflation. The suggested parachute types for the 2x2U LEO configuration are the ringslot, disk-gap-band, extended skirt 14.3% full, and conical ribbon parachutes, due to their low opening forces. The parachute should be deployed at a Mach number less than 0.65 to minimize opening decelerations. However, these designed parachutes do not consistently meet the impact deceleration requirement in a Monte Carlo simulation and should be oversized to account for variability. This strategy is applied to the 1U LEO configuration and results in approximately 49% confidence of meeting the impact velocity requirement with doubling the parachute area. Only 63% confidence is achieved by tripling the parachute area, indicating significantly diminished returns with increasing area. These approximate confidence levels are present with all eleven parachute types.
Dynamic Stability Analysis of Blunt Body Entry Vehicles Through the Use of a Time-Lagged Aftbody Pitching Moment

(Georgia Institute of Technology, 2012-10-05) Kazemba, Cole

This analysis defines an analytic model for the pitching motion of blunt bodies during atmospheric entry. The proposed model is independent of the pitch damping sum term which is present in the standard equations of motion, instead using the principle of a time-lagged aftbody moment as the forcing function for oscillation divergence. Four parameters, all with intuitive physical relevance, are introduced to fully define the aftbody moment and the associated time delay. It is shown that the dynamic oscillation responses typical to blunt bodies can be produced using hysteresis of the aftbody moment alone. The approach used in this investigation is shown to be useful in understanding the governing physical mechanisms for blunt body dynamic stability and in guiding vehicle and mission design requirements. A case study using simulated ballistic range test data is conducted. From this, parameter identification is carried out through the use of a least squares optimizing routine. Results show good agreement with the limited existing literature for the parameters identified. The model parameters were found to be accurate for a wide array of initial conditions and can be identified with a reasonable number of ballistic range shots and computational effort.
SPORE Mission Design

(Georgia Institute of Technology, 2012-08-01) Bauer, Nicole

Small Probes for Orbital Return of Experiments (SPORE), provides an on-orbit and re-entry platform for a range of biological, thermal protection system (TPS) characterization, and material science experiments. This platform will provide the capability for 1-4 weeks of on orbit flight operations for experiments with comparable mass and volumes laid out by the 1U and 2U cubesat guidelines. The platform will accommodate return from low earth orbit (LEO) and geosynchronous transfer orbit (GTO) to maximize the science potential. Packaging models and mass budgets were created for the SPORE entry vehicles. In addition, analyses were completed to construct the SPORE mission design. Orbital trajectory and maneuvers were modeled for the LEO and GTO missions. A re-entry architecture was designed to meet the requirements set by the range of payloads, orbits, and entry vehicle sizes inherent in the SPORE mission concept. The selected entry, descent, and landing (EDL) architecture was validated and modeled using 3-DOF and 6-DOF software. A thermal soak-back analysis was used to generate a temperature profile for the payload, and a Monte Carlo analysis was completed on this architecture to assess landing footprint. Results from the landing dispersion analysis confirm current landing site selections and help establish recovery procedures.
Guided Entry Performance of Low Ballistic Coefficient Vehicles at Mars

(Georgia Institute of Technology, 2012-05-21) Meginnis, Ian M.

Current Mars entry, descent, and landing technology is near its performance limit and is generally unable to land payloads on the surface that exceed approximately 1 metric ton. One option for increasing landed payload mass capability is decreasing the entry vehicle’s hypersonic ballistic coefficient. A lower ballistic coefficient vehicle decelerates higher in the atmosphere, providing additional timeline and altitude margin necessary for landing more massive payloads. This study analyzed the guided entry performance of several low ballistic coefficient vehicle concepts at Mars. A terminal point controller guidance algorithm, based on the Apollo Final Phase algorithm, was used to provide precision targeting capability. Terminal accuracy, peak deceleration, peak heat rate, and integrated heat load were assessed and compared to a traditional Mars entry vehicle concept to determine the effects of lowering the vehicle ballistic coefficient on entry performance. Results indicate that, while terminal accuracy degrades slightly with decreasing ballistic coefficient, the terminal accuracy and other performance metrics remain within reasonable bounds for ballistic coefficients as low as 1 kg/m2 . As such, this investigation demonstrates that from a performance standpoint, guided entry vehicles with low ballistic coefficients (large diameters) may be feasible at Mars. Additionally, flight performance may be improved through the use of guidance schemes designed specifically for low ballistic coefficient vehicles, as well as novel terminal descent systems designed around low ballistic coefficient trajectories
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
Automated Proximity Operations Using Image Based Relative Navigation

(Georgia Institute of Technology, 2012-05-01) Walker, Luke

This paper will describe a system for relative navigation and automated proximity operations for a small spacecraft about another spacecraft using continuous thrust propulsion and low cost imagers. Novel image processing algorithms provide range estimates in addition to traditional spherical angle estimates using knowledge of the target spacecraft’s geometry. A differential correction batch filter is used to provide relative navigation and state estimation. These state estimates are used to provide input for the automated control of the chaser spacecraft via a Linear Quadratic Regulator. Propulsive maneuvers are accomplished using several low-thrust, non-throttleable thrusters using pulse-width modulation and thrust vectoring. A waypoint logic controller is used to define intermediate goals to reach the final goal in order to limit operational risk from an error in estimation of the spacecraft’s relative state. The system is described and then initial simulation test results are shown.
Relative Navigation for Satellites in Close Proximity Using Angles-Only Observations

(Georgia Institute of Technology, 2011-12-01) Patel, Hemanshu

Relative navigation using angles-only observations is explored in this research. Previous work has shown that the unique relative orbit of a deputy satellite cannot be found using angles-only camera measurements from the chief satellite when a linear model of relative motion is used, due to a lack of observability. This work examines the possibility of partial observability in this case, which consists of a basis vector that corresponds to a family of relative orbits. A Preliminary Orbit Determination (POD) method is introduced that uses 3 Line-Of-Sight (LOS) measurements and provides an initial guess for the basis vector. This guess is differentially corrected with a batch filter that takes in a full set of LOS measurements to hone in on a converged solution for the basis vector. The application of an Extended Kalman Filter (EKF) to this problem is also explored.