You searched for subject:(Astrodynamics).
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Embry-Riddle Aeronautical University
1.
Iyengar, Spatika Dasharati.
Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model.
Degree: MSin Aerospace Engineering, Graduate Studies, 2015, Embry-Riddle Aeronautical University
URL: https://commons.erau.edu/edt/251 
► Thermal Protection systems (TPS) are used as shields in space vehicles which encounter high heat and temperatures at the reentry altitudes. Among them, the…
(more)
▼ Thermal Protection systems (TPS) are used as shields in space vehicles which encounter high heat and temperatures at the reentry altitudes. Among them, the cooling techniques and the ablative coatings are most popular. However, they have their own weight limitations. In the recent decade, another classification of TPS called the Non-Ablative Thermal Protection systems (NaTPS) have gained significance. The spike-lateral jet method is an NaTPS concept proposed for drag and heat flux reduction in hypersonic nose cones. Numerical simulations are conducted to analyze the effectiveness of spike-lateral jet concept at re-entry altitudes. The spike attached to the hemispherical nose has two circular orifices which eject air. The freestream conditions include Mach number 6 and standard atmospheric conditions at 30km altitude. The k-w turbulent model is used to model the case. It is apparent from the results that the lateral jet reconstructs the flow field by pushing the conical shock away and creating a large recirculation zone in front of the blunt body. This pushes the reattachment region rearward thus decreasing peak pressure and heat transfer to the body. The peak pressure at the flow reattachment point on the blunt body can be reduced by 35.9% for the laminar case and 30.0% for the turbulent case. The heat flux can be reduced by 54.1% for laminar case and 64.3% for the turbulent case. Lateral jet injection did not reduce drag as proposed, instead it increased drag by 65.7% and 59.1% for laminar and turbulent cases respectively. Base drag is not included in drag calculations. Results show that it might be promising for future applications of heat flux reduction and peak pressure reduction in reentry vehicles.
Subjects/Keywords: Astrodynamics
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APA (6th Edition):
Iyengar, S. D. (2015). Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model. (Masters Thesis). Embry-Riddle Aeronautical University. Retrieved from https://commons.erau.edu/edt/251
Chicago Manual of Style (16th Edition):
Iyengar, Spatika Dasharati. “Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model.” 2015. Masters Thesis, Embry-Riddle Aeronautical University. Accessed March 04, 2021.
https://commons.erau.edu/edt/251.
MLA Handbook (7th Edition):
Iyengar, Spatika Dasharati. “Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model.” 2015. Web. 04 Mar 2021.
Vancouver:
Iyengar SD. Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model. [Internet] [Masters thesis]. Embry-Riddle Aeronautical University; 2015. [cited 2021 Mar 04].
Available from: https://commons.erau.edu/edt/251.
Council of Science Editors:
Iyengar SD. Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model. [Masters Thesis]. Embry-Riddle Aeronautical University; 2015. Available from: https://commons.erau.edu/edt/251
2.
Vittaldev, V. (author).
The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.
Degree: 2010, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:198390b8-b81f-4a13-9c80-15befd3c0428
Astrodynamics and Space Missions
Aerospace Engineering
Advisors/Committee Members: Mooij, E. (mentor), Naeije, M.C. (mentor).Subjects/Keywords: Astrodynamics
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4 Astrodynamics and Orbital Mechanics
4.1 Gravity… …of astrodynamics and the space
environment with the various perturbations…
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APA (6th Edition):
Vittaldev, V. (. (2010). The Unified State Model. Derivation and Applications in Astrodynamics and Navigation. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:198390b8-b81f-4a13-9c80-15befd3c0428
Chicago Manual of Style (16th Edition):
Vittaldev, V (author). “The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.” 2010. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:198390b8-b81f-4a13-9c80-15befd3c0428.
MLA Handbook (7th Edition):
Vittaldev, V (author). “The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.” 2010. Web. 04 Mar 2021.
Vancouver:
Vittaldev V(. The Unified State Model. Derivation and Applications in Astrodynamics and Navigation. [Internet] [Masters thesis]. Delft University of Technology; 2010. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:198390b8-b81f-4a13-9c80-15befd3c0428.
Council of Science Editors:
Vittaldev V(. The Unified State Model. Derivation and Applications in Astrodynamics and Navigation. [Masters Thesis]. Delft University of Technology; 2010. Available from: http://resolver.tudelft.nl/uuid:198390b8-b81f-4a13-9c80-15befd3c0428
Embry-Riddle Aeronautical University
3.
Sestak, Timothy Allen.
The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics.
Degree: PhD, College of Aviation, 2017, Embry-Riddle Aeronautical University
URL: https://commons.erau.edu/edt/355
► This study examines the aerodynamic effects of the materials, textiles, and morphologies currently used in wingsuit design and construction. The experiment was a low-speed…
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▼ This study examines the aerodynamic effects of the materials, textiles, and morphologies currently used in wingsuit design and construction. The experiment was a low-speed wind tunnel investigation using a rigid wing with an aspect ratio of 2, a NACA 4418 airfoil cross section and a smooth, polished painted surface as a baseline. The baseline wing was modified by covering the upper and lower surfaces with various textiles currently used in wingsuit construction. This study is the first step in continued research to design and build a wingsuit with superior glide performance compared to current designs. Surface textures and features on the lifting surfaces of wings are known to have significant aerodynamic consequences. This experiment compared the lift and drag of a representative low aspect ratio wing before and after covering the wing with the various fabrics and textiles used in current wingsuit design and arranged the various textiles and other wingsuit features, like zippers and seams, in morphologies currently used in wingsuit construction.
The data collected clearly shows current wingsuit materials and morphologies have a potentially large, usually undesirable effect on flight performance. All woven fabrics reduced aerodynamic efficiency as measured by CL/CD. Those treatments with the roughest surfaces greatly reduced lift and increased drag as much as 50% or more and reduced aerodynamic efficiency as much as 75%. Placement of zippers and seams are shown to be critical factors for both aerodynamic efficiency and stability. Current combinations of fabrics and morphologies were shown to be often mutually and additively detrimental to aerodynamic performance. Certain textiles showed possible utility in drag reduction. It was initially thought that the effects of the surface treatments on lift would be the major factor in wingsuit performance. While the effects on lift were significant, the large drag penalties due to woven and textured fabrics and textiles and the early separation of airflow at low angles of attack, appear to have had the greatest effect on the aerodynamic efficiency of a lifting surface with an airfoil cross section.
Subjects/Keywords: Astrodynamics; Aviation
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APA (6th Edition):
Sestak, T. A. (2017). The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics. (Doctoral Dissertation). Embry-Riddle Aeronautical University. Retrieved from https://commons.erau.edu/edt/355
Chicago Manual of Style (16th Edition):
Sestak, Timothy Allen. “The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics.” 2017. Doctoral Dissertation, Embry-Riddle Aeronautical University. Accessed March 04, 2021.
https://commons.erau.edu/edt/355.
MLA Handbook (7th Edition):
Sestak, Timothy Allen. “The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics.” 2017. Web. 04 Mar 2021.
Vancouver:
Sestak TA. The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics. [Internet] [Doctoral dissertation]. Embry-Riddle Aeronautical University; 2017. [cited 2021 Mar 04].
Available from: https://commons.erau.edu/edt/355.
Council of Science Editors:
Sestak TA. The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics. [Doctoral Dissertation]. Embry-Riddle Aeronautical University; 2017. Available from: https://commons.erau.edu/edt/355
University of Florida
4.
Kelly, Patrick W.
Orbital Debris Mitigation Using Solar Sail Trajectory Optimization.
Degree: PhD, Aerospace Engineering - Mechanical and Aerospace Engineering, 2018, University of Florida
URL: https://ufdc.ufl.edu/UFE0053995
► Solar-sailing satellites offer unique propulsive advantages for orbit manipulation and propellant mass requirements. When compared with traditional, propellant-based satellites, solar-sailing spacecraft are not limited by…
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▼ Solar-sailing satellites offer unique propulsive advantages for orbit manipulation and propellant mass requirements. When compared with traditional, propellant-based satellites, solar-sailing spacecraft are not limited by propellant consumption or Δ V availability. By harnessing the existing elements of a star-based system, solar-sails take advantage of ``free'' propulsion by means of momentum exchange with incoming photons. Given the current trend of satellite miniaturization, solar-sails may provide a feasible propulsive solution for small satellites in the near future.
Advisors/Committee Members: BEVILACQUA,RICCARDO (committee chair), RAO,ANIL (committee member), FRY,JAMES N (committee member).
Subjects/Keywords: astrodynamics – optimization
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APA (6th Edition):
Kelly, P. W. (2018). Orbital Debris Mitigation Using Solar Sail Trajectory Optimization. (Doctoral Dissertation). University of Florida. Retrieved from https://ufdc.ufl.edu/UFE0053995
Chicago Manual of Style (16th Edition):
Kelly, Patrick W. “Orbital Debris Mitigation Using Solar Sail Trajectory Optimization.” 2018. Doctoral Dissertation, University of Florida. Accessed March 04, 2021.
https://ufdc.ufl.edu/UFE0053995.
MLA Handbook (7th Edition):
Kelly, Patrick W. “Orbital Debris Mitigation Using Solar Sail Trajectory Optimization.” 2018. Web. 04 Mar 2021.
Vancouver:
Kelly PW. Orbital Debris Mitigation Using Solar Sail Trajectory Optimization. [Internet] [Doctoral dissertation]. University of Florida; 2018. [cited 2021 Mar 04].
Available from: https://ufdc.ufl.edu/UFE0053995.
Council of Science Editors:
Kelly PW. Orbital Debris Mitigation Using Solar Sail Trajectory Optimization. [Doctoral Dissertation]. University of Florida; 2018. Available from: https://ufdc.ufl.edu/UFE0053995
Cal Poly
5.
Farahmand, Mitra.
ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK.
Degree: MS, Aerospace Engineering, 2009, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/167
;
10.15368/theses.2009.138
► This thesis describes the models of four common orbital propagators and outlines the process of integrating them into the Horizon Simulation Framework (HSF). The results…
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▼ This thesis describes the models of four common orbital propagators and outlines the process of integrating them into the Horizon Simulation Framework (HSF). The results of the Two-Body, J2, and J4 propagators from the HSF are then compared against the outcomes of these propagators in MATLAB and Satellite Toolkit (STK). The MATLAB algorithms verify the functionality of the propagators and determine the accuracy of the HSF implementation. The compassion against STK validates the formulation of the HSF propagators.
In order to equip the HSF with a more precise means of orbit determination, adding the Simplified General Perturbations 4 (SGP4) propagator to the HSF has been the principal goal of this project. A brief description of the algorithm explains the process of configuring the original code into a format compatible with the HSF. Further, the orbital data from the SGP4 propagator across different implementations are examined. The outcomes demonstrate that the HSF algorithm generates reasonably accurate orbital data.
Advisors/Committee Members: Eric A. Mehiel.
Subjects/Keywords: Astrodynamics
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APA (6th Edition):
Farahmand, M. (2009). ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/167 ; 10.15368/theses.2009.138
Chicago Manual of Style (16th Edition):
Farahmand, Mitra. “ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK.” 2009. Masters Thesis, Cal Poly. Accessed March 04, 2021.
https://digitalcommons.calpoly.edu/theses/167 ; 10.15368/theses.2009.138.
MLA Handbook (7th Edition):
Farahmand, Mitra. “ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK.” 2009. Web. 04 Mar 2021.
Vancouver:
Farahmand M. ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK. [Internet] [Masters thesis]. Cal Poly; 2009. [cited 2021 Mar 04].
Available from: https://digitalcommons.calpoly.edu/theses/167 ; 10.15368/theses.2009.138.
Council of Science Editors:
Farahmand M. ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK. [Masters Thesis]. Cal Poly; 2009. Available from: https://digitalcommons.calpoly.edu/theses/167 ; 10.15368/theses.2009.138
Purdue University
6.
Haapala, Amanda F.
Trajectory design in the spatial circular restricted three-body problem exploiting higher-dimensional Poincare maps.
Degree: PhD, Aeronautics and Astronautics, 2014, Purdue University
URL: https://docs.lib.purdue.edu/open_access_dissertations/278
► In this investigation, the role of higher-dimensional Poincaré maps in facilitating trajectory design is explored for a variety of applications. To begin, existing strategies…
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▼ In this investigation, the role of higher-dimensional Poincaré maps in facilitating trajectory design is explored for a variety of applications. To begin, existing strategies to implement Poincaré maps for trajectory design applications in the spatial CR3BP are evaluated. New applications for these strategies are explored, including an analysis of the natural motion of Jupiter-family comets that experience temporary capture about Jupiter, and the search for periodic orbits in the vicinity of the primary bodies in the spatial problem. Because current strategies to represent higher-dimensional maps, generally, lead to a loss of information, new approaches to represent all information contained in higher-dimensional Poincaré maps are sought. ^ The field of data visualization offers many options to visually represent multivariate data sets, including the use of glyphs. A glyph is any graphical object whose physical attributes are determined by the variables of a data set. In this investigation, the role of glyphs in representing higher-dimensional Poincaré maps is explored, and the resulting map representations are demonstrated to search for maneuver-free and low-cost transfers between libration point orbits. A catalog of libration point orbit transfers is developed in the Earth-Moon system, and observations about the catalog solutions yields insight into the existence of these transfers. The application of Poincaré maps to compute transfers between libration point orbits in different three-body systems is additionally considered. Finally, interactive trajectory design environments that incorporate Poincaré maps into the design process are demonstrated. Such design environments offer a unique opportunity to explore the available trajectory options and to gain intuition about the solution space.
Advisors/Committee Members: Kathleen C. Howell, Kathleen Howell, James M. Longuski, Martin J. Corless, Belinda Marchand.
Subjects/Keywords: Aerospace Engineering; Astrodynamics
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APA (6th Edition):
Haapala, A. F. (2014). Trajectory design in the spatial circular restricted three-body problem exploiting higher-dimensional Poincare maps. (Doctoral Dissertation). Purdue University. Retrieved from https://docs.lib.purdue.edu/open_access_dissertations/278
Chicago Manual of Style (16th Edition):
Haapala, Amanda F. “Trajectory design in the spatial circular restricted three-body problem exploiting higher-dimensional Poincare maps.” 2014. Doctoral Dissertation, Purdue University. Accessed March 04, 2021.
https://docs.lib.purdue.edu/open_access_dissertations/278.
MLA Handbook (7th Edition):
Haapala, Amanda F. “Trajectory design in the spatial circular restricted three-body problem exploiting higher-dimensional Poincare maps.” 2014. Web. 04 Mar 2021.
Vancouver:
Haapala AF. Trajectory design in the spatial circular restricted three-body problem exploiting higher-dimensional Poincare maps. [Internet] [Doctoral dissertation]. Purdue University; 2014. [cited 2021 Mar 04].
Available from: https://docs.lib.purdue.edu/open_access_dissertations/278.
Council of Science Editors:
Haapala AF. Trajectory design in the spatial circular restricted three-body problem exploiting higher-dimensional Poincare maps. [Doctoral Dissertation]. Purdue University; 2014. Available from: https://docs.lib.purdue.edu/open_access_dissertations/278
Penn State University
7.
Conte, Davide.
Determination of Optimal Earth-mars Trajectories to Target the Moons of Mars.
Degree: 2014, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/21382
► The focus of this thesis is to analyze interplanetary transfer maneuvers from Earth to Mars in order to target the Martian moons, Phobos and Deimos.…
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▼ The focus of this thesis is to analyze interplanetary transfer maneuvers from Earth to Mars in order to target the Martian moons, Phobos and Deimos. Such analysis is done by solving Lambert’s Problem and investigating the necessary targeting upon Mars arrival. Additionally, the orbital parameters of the arrival trajectory as well as the relative required ΔVs and times of flights were determined in order to define the optimal departure and arrival windows for a given range of date. The first step in solving Lambert’s Problem consists in finding the positions and velocities of the departure (Earth) and arrival (Mars) planets for a given range of dates. Then, by solving Lambert’s problem for various combinations of departure and arrival dates, porkchop plots can be created and examined. Some of the key parameters that are plotted on porkchop plots and used to investigate possible transfer orbits are the departure characteristic energy, C3, and the arrival v∞. These parameters, in combination with given desired initial and final parking orbital conditions about Earth and Mars, were also used to determine the total ΔV for the various Earth-Mars transfers. ΔV results were used to find the necessary amount of propellant needed for the transfer maneuvers as a percentage of total spacecraft mass for a given specific impulse. Moreover, this thesis considers the arrival trajectories of an arbitrary spacecraft with respect to a Mars inertial reference frame. Key parameters are the inclination of the arrival orbit with respect to the Mars equator and ΔV needed to be captured by Mars and be inserted in the Martian moons’ orbital planes. Lastly, rendezvous maneuvers with Phobos/Deimos are considered.
The results of the analysis lead to the optimal choice of departure and arrival dates that, given the capabilities of the launch system such as ΔV, makes the mission physically feasible.
Advisors/Committee Members: Dr David Spencer, Thesis Advisor/Co-Advisor.
Subjects/Keywords: Mars; Phobos; Deimos; astrodynamics
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APA (6th Edition):
Conte, D. (2014). Determination of Optimal Earth-mars Trajectories to Target the Moons of Mars. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/21382
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Conte, Davide. “Determination of Optimal Earth-mars Trajectories to Target the Moons of Mars.” 2014. Thesis, Penn State University. Accessed March 04, 2021.
https://submit-etda.libraries.psu.edu/catalog/21382.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Conte, Davide. “Determination of Optimal Earth-mars Trajectories to Target the Moons of Mars.” 2014. Web. 04 Mar 2021.
Vancouver:
Conte D. Determination of Optimal Earth-mars Trajectories to Target the Moons of Mars. [Internet] [Thesis]. Penn State University; 2014. [cited 2021 Mar 04].
Available from: https://submit-etda.libraries.psu.edu/catalog/21382.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Conte D. Determination of Optimal Earth-mars Trajectories to Target the Moons of Mars. [Thesis]. Penn State University; 2014. Available from: https://submit-etda.libraries.psu.edu/catalog/21382
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Embry-Riddle Aeronautical University
8.
Coulter, Nolan.
Design of an Attitude Control System for a Spacecraft with Propellant Slosh Dynamics.
Degree: MSin Aerospace Engineering, Graduate Studies, 2018, Embry-Riddle Aeronautical University
URL: https://commons.erau.edu/edt/424
► The presence of propellant slosh dynamics in a spacecraft system during a maneuver leads to attitude control system (ACS) performance degradation resulting in attitude…
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▼ The presence of propellant slosh dynamics in a spacecraft system during a maneuver leads to attitude control system (ACS) performance degradation resulting in attitude tracking errors and instability. As spacecraft missions become more complex and involve longer durations, a substantial propellant mass is required to achieve the mission objectives and perform orbital maneuvers. When the propellant tanks are only partially filled, the liquid fuel moves inside the tanks with translational and rotational accelerations generating the slosh dynamics. This research effort performs a comparative study with different optimal control techniques and a novel application of a model reference artificial immune system adaptive controller (MRAIS). A linearized model of a realistic spacecraft dynamic model incorporating propellant slosh is derived utilizing the mass-spring analogy. Simulations with the linearized models assist in control law development to achieve the control objective: to suppress the fuel slosh dynamics while obtaining the desired attitude. These control laws are then tested with the nonlinear equations of motion for a spacecraft with propellant slosh dynamics to evaluate the ability of the models to design an attitude control system. Monte Carlo analysis is also applied to characterize the performance of each controller and determine the most significant parameters that cause instability issues. The Model Reference Artificial Immune System has superior performance in comparison to the baseline optimal control systems and is more robust to system instabilities, actuator failures, and aggressive maneuvers.
Subjects/Keywords: Aerospace Engineering; Astrodynamics; Space Vehicles
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APA ·
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APA (6th Edition):
Coulter, N. (2018). Design of an Attitude Control System for a Spacecraft with Propellant Slosh Dynamics. (Masters Thesis). Embry-Riddle Aeronautical University. Retrieved from https://commons.erau.edu/edt/424
Chicago Manual of Style (16th Edition):
Coulter, Nolan. “Design of an Attitude Control System for a Spacecraft with Propellant Slosh Dynamics.” 2018. Masters Thesis, Embry-Riddle Aeronautical University. Accessed March 04, 2021.
https://commons.erau.edu/edt/424.
MLA Handbook (7th Edition):
Coulter, Nolan. “Design of an Attitude Control System for a Spacecraft with Propellant Slosh Dynamics.” 2018. Web. 04 Mar 2021.
Vancouver:
Coulter N. Design of an Attitude Control System for a Spacecraft with Propellant Slosh Dynamics. [Internet] [Masters thesis]. Embry-Riddle Aeronautical University; 2018. [cited 2021 Mar 04].
Available from: https://commons.erau.edu/edt/424.
Council of Science Editors:
Coulter N. Design of an Attitude Control System for a Spacecraft with Propellant Slosh Dynamics. [Masters Thesis]. Embry-Riddle Aeronautical University; 2018. Available from: https://commons.erau.edu/edt/424
Delft University of Technology
9.
Agrawal, Abhishek (author).
Orbital motion of regolith around asteroids.
Degree: 2018, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:0933c1e6-89ee-476c-9b45-3e6d235ddac6
► We study the orbital motion of regolith around asteroids, lofted from the surface due to impact cratering events, to understand the displacement of material on…
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▼ We study the orbital motion of regolith around asteroids, lofted from the surface due to impact cratering events, to understand the displacement of material on the surface and in orbit. The cratering events could be natural such as from meteoroid impacts, or they can be induced from spacecraft activities such as in-situ sample collection. Understanding the dynamics of orbiting regolith is important for future science missions and commercial activities on asteroids. Knowledge about expected particulate environment due to impact ejecta can help mission designers in trajectory planning to avoid interference or damage from orbiting regolith with a spacecraft and/or its instruments. The same study could be exploited in the field of commercial in-situ asteroid mining for sorting material of different sizes and densities by artificially lofting them into an orbit...
Advisors/Committee Members: Noomen, Ron (mentor), Visser, Pieter (graduation committee), Cervone, Angelo (graduation committee), Delft University of Technology (degree granting institution).
Subjects/Keywords: Astrodynamics; Regolith; Asteroids; Numerical Simulation
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APA (6th Edition):
Agrawal, A. (. (2018). Orbital motion of regolith around asteroids. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:0933c1e6-89ee-476c-9b45-3e6d235ddac6
Chicago Manual of Style (16th Edition):
Agrawal, Abhishek (author). “Orbital motion of regolith around asteroids.” 2018. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:0933c1e6-89ee-476c-9b45-3e6d235ddac6.
MLA Handbook (7th Edition):
Agrawal, Abhishek (author). “Orbital motion of regolith around asteroids.” 2018. Web. 04 Mar 2021.
Vancouver:
Agrawal A(. Orbital motion of regolith around asteroids. [Internet] [Masters thesis]. Delft University of Technology; 2018. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:0933c1e6-89ee-476c-9b45-3e6d235ddac6.
Council of Science Editors:
Agrawal A(. Orbital motion of regolith around asteroids. [Masters Thesis]. Delft University of Technology; 2018. Available from: http://resolver.tudelft.nl/uuid:0933c1e6-89ee-476c-9b45-3e6d235ddac6
Delft University of Technology
10.
de Jong, Hanneke (author).
Analytical Low-Thrust Trajectory Design: Using the Simplified General Perturbations Model.
Degree: 2018, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:1c58ca05-b367-48e6-9b4c-e31f12f488d7
► An analytical low-thrust design tool for orbits around Earth has been developed which takes perturbations into account. The Simplified General Perturbations 4 model (SGP4) is…
(more)
▼ An analytical low-thrust design tool for orbits around Earth has been developed which takes perturbations into account. The Simplified General Perturbations 4 model (SGP4) is combined with Edelbaum’s analytical low-thrust trajectory model. This resulted in the SGP4-LT tool which required an iterative version of SGP4. To obtain this iterative SGP4, an existing method was corrected and extended. A convergence rate of 99.5% was obtained for 17542 objects of the satellite catalogue. Convergence was not reached for combinations of inclinations smaller than 0.1° and eccentricities lower than 4e-6. SGP4-LT obtained expected results for orbit raising and non-coplanar orbit change. Altitude maintenance was performed which was based on the GOCE mission. SGP4-LT obtained a 14 % higher amount of propellant than GOCE had available. SGP4-LT provides within seconds a first approximation solution for low-thrust transfer trajectories around Earth and is capable of calculating DeltaV budgets for altitude maintenance in low Earth orbits.
Aerospace Engineering
Advisors/Committee Members: Noomen, Ron (mentor), Delft University of Technology (degree granting institution).
Subjects/Keywords: astrodynamics; SGP4; Low-thrust
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APA (6th Edition):
de Jong, H. (. (2018). Analytical Low-Thrust Trajectory Design: Using the Simplified General Perturbations Model. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:1c58ca05-b367-48e6-9b4c-e31f12f488d7
Chicago Manual of Style (16th Edition):
de Jong, Hanneke (author). “Analytical Low-Thrust Trajectory Design: Using the Simplified General Perturbations Model.” 2018. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:1c58ca05-b367-48e6-9b4c-e31f12f488d7.
MLA Handbook (7th Edition):
de Jong, Hanneke (author). “Analytical Low-Thrust Trajectory Design: Using the Simplified General Perturbations Model.” 2018. Web. 04 Mar 2021.
Vancouver:
de Jong H(. Analytical Low-Thrust Trajectory Design: Using the Simplified General Perturbations Model. [Internet] [Masters thesis]. Delft University of Technology; 2018. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:1c58ca05-b367-48e6-9b4c-e31f12f488d7.
Council of Science Editors:
de Jong H(. Analytical Low-Thrust Trajectory Design: Using the Simplified General Perturbations Model. [Masters Thesis]. Delft University of Technology; 2018. Available from: http://resolver.tudelft.nl/uuid:1c58ca05-b367-48e6-9b4c-e31f12f488d7
University of Illinois – Urbana-Champaign
11.
Ponnapalli, Kaushik.
Optimal variable-specific-impulse trajectories in simple gravitational fields.
Degree: MS, Aerospace Engineering, 2019, University of Illinois – Urbana-Champaign
URL: http://hdl.handle.net/2142/104920
► Minimum-fuel trajectories in the Clohessy-Wiltshire orbital model are investigated. Low-thrust variable-specific-impulse propulsion is assumed. The necessary conditions for an optimal solution and the equations of…
(more)
▼ Minimum-fuel trajectories in the Clohessy-Wiltshire orbital model are investigated. Low-thrust variable-specific-impulse propulsion is assumed. The necessary conditions for an optimal solution and the equations of motion are combined to produce a single equation. Every solution to this equation is an optimal trajectory. Boundary conditions for rendezvous and intercept trajectories are formulated and optimal solutions are obtained. Previous analytical solutions are examined and reconciled.
Advisors/Committee Members: Prussing, John E (advisor).
Subjects/Keywords: Astrodynamics, Optimal Trajectories; Trajectory Optimization
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APA (6th Edition):
Ponnapalli, K. (2019). Optimal variable-specific-impulse trajectories in simple gravitational fields. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/104920
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Ponnapalli, Kaushik. “Optimal variable-specific-impulse trajectories in simple gravitational fields.” 2019. Thesis, University of Illinois – Urbana-Champaign. Accessed March 04, 2021.
http://hdl.handle.net/2142/104920.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Ponnapalli, Kaushik. “Optimal variable-specific-impulse trajectories in simple gravitational fields.” 2019. Web. 04 Mar 2021.
Vancouver:
Ponnapalli K. Optimal variable-specific-impulse trajectories in simple gravitational fields. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2019. [cited 2021 Mar 04].
Available from: http://hdl.handle.net/2142/104920.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Ponnapalli K. Optimal variable-specific-impulse trajectories in simple gravitational fields. [Thesis]. University of Illinois – Urbana-Champaign; 2019. Available from: http://hdl.handle.net/2142/104920
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of New South Wales
12.
Crowe, William.
Use of Spacecraft Swarms to Characterise Asteroids during Flyby.
Degree: Mechanical & Manufacturing Engineering, 2018, University of New South Wales
URL: http://handle.unsw.edu.au/1959.4/60892
► Of the more than 700,000 asteroids known today, only 12 have been visited by spacecraft. This means that many assumptions about their characteristics remain unverified.…
(more)
▼ Of the more than 700,000 asteroids known today, only 12 have been visited by spacecraft. This means that many assumptions about their characteristics remain unverified. Various swarms of spacecraft have been suggested for characterising asteroids. The development of these spacecraft swarms could improve the versatility of these missions and increase the number of asteroids visited. In this thesis, two novel swarm flyby strategies are created that build upon previous spacecraft swarm development. The first of these strategies uses the relative dynamics of a swarm to estimate an asteroid’s mass. Analytic equations are created and verified using numerical analysis for this estimation. It is found that this technique can be implemented for asteroids more massive than 1×〖10〗
18 kilograms, roughly the size of 21 Lutetia. A swarm of three or more spacecraft can also allow for the passing distance of spacecraft from the asteroid to be estimated simultaneously. However, this can result in errors greater than 10% if flyby conditions are not carefully controlled. It is also found that the relative dynamics between swarm spacecraft allow for the absolute swarm orientation to be determined, although this option is unattractive when mission operations are considered.The second strategy sees a swarm of spacecraft in orbit around the Earth waiting for an asteroid to pass close before moving out to meet it. Data from 2016 is used to find how many were accessible to the swarm using analytic orbit analysis. The probability that a spacecraft would lie at a position on the holding orbit where it could reach an asteroid for a given ΔV is used to make a decision on whether to transfer. From this data, up to five asteroids could be visited from High Earth Orbit with a ΔV of less than 1 km/s per asteroid. More asteroids are likely to become accessible using this method if new observation techniques are employed.The use of swarm spacecraft to estimate asteroid mass reduces dependence on Deep Space Network antennas. A swarm can make its estimation using a limited number of sensors, although they must be repeated across several spacecraft. Finally, a large number of asteroids can be visited each year if a swarm of spacecraft is placed in Earth orbit and used to visit close-passing asteroids.
Advisors/Committee Members: Olsen, John, Mechanical & Manufacturing Engineering, Faculty of Engineering, UNSW, Page, John, Mechanical & Manufacturing Engineering, Faculty of Engineering, UNSW.
Subjects/Keywords: Spacecraft; Asteroids; Astrodynamics; Swarms
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APA (6th Edition):
Crowe, W. (2018). Use of Spacecraft Swarms to Characterise Asteroids during Flyby. (Doctoral Dissertation). University of New South Wales. Retrieved from http://handle.unsw.edu.au/1959.4/60892
Chicago Manual of Style (16th Edition):
Crowe, William. “Use of Spacecraft Swarms to Characterise Asteroids during Flyby.” 2018. Doctoral Dissertation, University of New South Wales. Accessed March 04, 2021.
http://handle.unsw.edu.au/1959.4/60892.
MLA Handbook (7th Edition):
Crowe, William. “Use of Spacecraft Swarms to Characterise Asteroids during Flyby.” 2018. Web. 04 Mar 2021.
Vancouver:
Crowe W. Use of Spacecraft Swarms to Characterise Asteroids during Flyby. [Internet] [Doctoral dissertation]. University of New South Wales; 2018. [cited 2021 Mar 04].
Available from: http://handle.unsw.edu.au/1959.4/60892.
Council of Science Editors:
Crowe W. Use of Spacecraft Swarms to Characterise Asteroids during Flyby. [Doctoral Dissertation]. University of New South Wales; 2018. Available from: http://handle.unsw.edu.au/1959.4/60892
Delft University of Technology
13.
Vandenberghe, Alexander (author).
Station-keeping in the vicinity of the collinear Lagrange points.
Degree: 2020, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:8b82e473-2ed1-48c4-ac3e-a59ee1a5039c
► In this thesis, the performance of three different station-keeping methods applied to Sun-Earth L2 halo orbits is investigated. In order to evaluate these methods, first…
(more)
▼ In this thesis, the performance of three different station-keeping methods applied to Sun-Earth L2 halo orbits is investigated. In order to evaluate these methods, first a framework for accurate satellite dynamics was developed. This was done by starting off at the Circular Restricted Three Body Problem, and extending that to a more accurate model in the Roto-Pulsating Reference Frame (RPF). The dynamics in the RPF were implemented in a propagator, written in C, which forms the basis for any analysis that has been done for this thesis. The propagator includes real planetary ephemerides from JPL, perturbations from all solar-system bodies and Solar Radiation Pressure and has functionalities for orbit generation, STM propagation, reference frame transformations and manoeuvre calculations. The three station-keeping methods that have been investigated in this thesis are: Optimal Continuation Strategy (OCS), Discrete-time Sliding Mode Control (DSMC) and Fuzzy-logic Sliding Mode Control (FLSMC). All of these methods were implemented in C, within the developed propagator and were tuned using an optimization library for C. After the implementation of the three methods, a performance analysis was done, applying the methods to a Sun-Earth L2 halo orbit in a Monte-Carlo simulation, taking into account uncertainties in the orbit determination process, the perturbations and the manoeuvre execution. The total ΔV, as well as the number of manoeuvres, for the three methods were calculated, and subsequently compared against eachother. This was done for a time span of 500 days with a minimum manoeuvre interval of 12 days. The resulting comparison and the result of this analysis clearly shows that the DSMC and FLSMC perform much better both in terms of total ΔV as well as in the variance that arises from the Monte-Carlo simulations. The FLSMC provides another significant, yet smaller, improvement when compared to the DSMC, which was theorized by Lian et al. The fuzzy-logic layer that is implemented makes for smoother behaviour when the actual orbit of the satellite is close to the desired reference orbit, where the normal DSMC exhibits more chattering around the reference orbit. It is clear from this analysis that the proposal of Lian et al. to add this fuzzy logic layer does contribute significantly to the performance of the method and that FLSMC is a potent and efficient method for halo orbit station-keeping, provided a good reference orbit is chosen for the method to follow. The analysis performed in this thesis was restricted to Sun-Earth L2 halo orbits but all techniques mentioned and developed within this research can easily be applied to halo orbits around any of the collinear lagrange points in any three-body system, with only minor changes.
Aerospace Engineering
Advisors/Committee Members: Noomen, R. (mentor), Herrera, Javier (graduation committee), Delft University of Technology (degree granting institution).
Subjects/Keywords: Station-keeping; Halo; Astrodynamics; Optimisation
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APA (6th Edition):
Vandenberghe, A. (. (2020). Station-keeping in the vicinity of the collinear Lagrange points. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:8b82e473-2ed1-48c4-ac3e-a59ee1a5039c
Chicago Manual of Style (16th Edition):
Vandenberghe, Alexander (author). “Station-keeping in the vicinity of the collinear Lagrange points.” 2020. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:8b82e473-2ed1-48c4-ac3e-a59ee1a5039c.
MLA Handbook (7th Edition):
Vandenberghe, Alexander (author). “Station-keeping in the vicinity of the collinear Lagrange points.” 2020. Web. 04 Mar 2021.
Vancouver:
Vandenberghe A(. Station-keeping in the vicinity of the collinear Lagrange points. [Internet] [Masters thesis]. Delft University of Technology; 2020. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:8b82e473-2ed1-48c4-ac3e-a59ee1a5039c.
Council of Science Editors:
Vandenberghe A(. Station-keeping in the vicinity of the collinear Lagrange points. [Masters Thesis]. Delft University of Technology; 2020. Available from: http://resolver.tudelft.nl/uuid:8b82e473-2ed1-48c4-ac3e-a59ee1a5039c
Cal Poly
14.
Graef, Jared.
B-Plane Targeting with the Spacecraft Trajectory Optimization Suite.
Degree: MS, Aerospace Engineering, 2020, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/2251
► In interplanetary trajectory applications, it is common to design arrival trajectories based on B-plane target values. This targeting scheme, B-plane targeting, allows for specific…
(more)
▼ In interplanetary trajectory applications, it is common to design arrival trajectories based on B-plane target values. This targeting scheme, B-plane targeting, allows for specific target orbits to be obtained during mission design. A primary objective of this work was to implement B-plane targeting into the Spacecraft Trajectory Optimization Suite (STOpS). This work was based on the previous versions of STOpS done by Fitzgerald and Sheehan, however STOpS was redeveloped from MATLAB to python. This updated version of STOpS implements 3-dimensional computation, departure and arrival orbital phase modeling with patched conics, B-plane targeting, and a trajectory correction maneuver. The optimization process is done with three evolutionary algorithms implemented in an island model paradigm.
The algorithms and the island model were successfully verified with known optimization functions before being used in the orbital optimization cases. While the algorithms and island model are not new to this work, they were altered in this redevelopment of STOpS to closer relate to literature. This enhanced literature relation allows for easier comprehension of the both the formulation of the schemes and the code itself. With a validated optimization scheme, STOpS is able to compute near-optimal trajectories for numerous historical missions. New mission types were also easily implemented and modeled with STOpS. A trajectory correction maneuver was shown to further optimize the trajectories end conditions, when convergence was reached. The result is a versatile optimization scheme that is highly customization to the invested user, while remaining simple for novice users.
Advisors/Committee Members: Kira Abercromby.
Subjects/Keywords: Optimization; Astrodynamics; B-Plane Targeting; Interplanetary; Astrodynamics; Space Vehicles
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APA (6th Edition):
Graef, J. (2020). B-Plane Targeting with the Spacecraft Trajectory Optimization Suite. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/2251
Chicago Manual of Style (16th Edition):
Graef, Jared. “B-Plane Targeting with the Spacecraft Trajectory Optimization Suite.” 2020. Masters Thesis, Cal Poly. Accessed March 04, 2021.
https://digitalcommons.calpoly.edu/theses/2251.
MLA Handbook (7th Edition):
Graef, Jared. “B-Plane Targeting with the Spacecraft Trajectory Optimization Suite.” 2020. Web. 04 Mar 2021.
Vancouver:
Graef J. B-Plane Targeting with the Spacecraft Trajectory Optimization Suite. [Internet] [Masters thesis]. Cal Poly; 2020. [cited 2021 Mar 04].
Available from: https://digitalcommons.calpoly.edu/theses/2251.
Council of Science Editors:
Graef J. B-Plane Targeting with the Spacecraft Trajectory Optimization Suite. [Masters Thesis]. Cal Poly; 2020. Available from: https://digitalcommons.calpoly.edu/theses/2251
University of Colorado
15.
Geeraert, Jeroen L.
Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications.
Degree: PhD, 2017, University of Colorado
URL: https://scholar.colorado.edu/asen_gradetds/192
► Very long baseline interferometry (VLBI) specifically same-beam interferometry (SBI), and dual-satellite geolocation are two fields of research not previously connected. This is due to…
(more)
▼ Very long baseline interferometry (VLBI) specifically same-beam interferometry (SBI), and dual-satellite geolocation are two fields of research not previously connected. This is due to the different application of each field, SBI is used for relative interplanetary navigation of two satellites while dual-satellite geolocation is used to locate the source of a radio frequency (RF) signal. In this dissertation however, we leverage both fields to create a novel method for multi-satellite orbit determination (OD) using time difference of arrival (TDOA) and frequency difference of arrival (FDOA) measurements. The measurements are double differenced between the satellites and the stations, in so doing, many of the common errors are canceled which can significantly improve measurement precision.
Provided with this novel OD technique, the observability is first analyzed to determine the benefits and limitations of this method. In all but a few scenarios the measurements successfully reduce the covariance when examining the Cramér-Rao Lower Bound (CRLB). Reduced observability is encountered with geostationary satellites as their motion with respect to the stations is limited, especially when only one baseline is used. However, when using satellite pairs with greater relative motion with respect to the stations, even satellites that are close to, but not exactly in a geostationary orbit can be estimated accurately. We find that in a strong majority of cases the OD technique provides lower uncertainties and solutions far more accurate than using conventional OD observables such as range and range-rate while also not being affected by common errors and biases. We specifically examine GEO-GEO, GEO-MEO, and GEO-LEO dual-satellite estimation cases. The work is further extended by developing a relative navigation scenario where the chief satellite is assumed to have perfect knowledge, or some small amount of uncertainty considered but not estimated, while estimating the deputy satellite state with respect to the chief. Once again the results demonstrate that the TDOA and FDOA OD results are favorable with faster dynamics over classical measurements.
This dissertation not only explores the OD side, but also gaps in geolocation research. First the mapping of ephemeris uncertainty to the geolocation covariance to provide a more realistic covariance was implemented. Furthermore, the geolocation solution was improved by appending a probabilistic altitude constraint to the posterior covariance, significantly reducing the projected geolocation uncertainty ellipse. The feasibility of using the geolocation setup to passively locate a LEO satellite was also considered. Finally the simulated results were verified using a long-arc of real data. The use of FDOA for small-body navigation and gravity recovery was also examined as an extended application.
Advisors/Committee Members: Jay W. McMahon, Penina Axelrad, Brandon Jones, Daniel Scheeres, Behrouz Touri.
Subjects/Keywords: astrodynamics; geolocation; interferometry; orbit determination; Aerospace Engineering
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APA (6th Edition):
Geeraert, J. L. (2017). Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications. (Doctoral Dissertation). University of Colorado. Retrieved from https://scholar.colorado.edu/asen_gradetds/192
Chicago Manual of Style (16th Edition):
Geeraert, Jeroen L. “Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications.” 2017. Doctoral Dissertation, University of Colorado. Accessed March 04, 2021.
https://scholar.colorado.edu/asen_gradetds/192.
MLA Handbook (7th Edition):
Geeraert, Jeroen L. “Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications.” 2017. Web. 04 Mar 2021.
Vancouver:
Geeraert JL. Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications. [Internet] [Doctoral dissertation]. University of Colorado; 2017. [cited 2021 Mar 04].
Available from: https://scholar.colorado.edu/asen_gradetds/192.
Council of Science Editors:
Geeraert JL. Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications. [Doctoral Dissertation]. University of Colorado; 2017. Available from: https://scholar.colorado.edu/asen_gradetds/192
Delft University of Technology
16.
Villegas Pinto, Daniel (author).
Temporary Capture of Asteroid Ejecta into Periodic Orbits: Application to JAXA’s Hayabusa2 Impact Event.
Degree: 2019, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:d9768a7f-0c57-4437-befb-4ccf7a39a0a5
► In the framework of JAXA’s Hayabusa2 mission, we study the dynamical environment around asteroid Ryugu to investigate whether ejecta particles can be temporarily trapped in…
(more)
▼ In the framework of JAXA’s Hayabusa2 mission, we study the dynamical environment around asteroid Ryugu to investigate whether ejecta particles can be temporarily trapped in periodic orbits following the Small Carry-on Impactor (SCI) operation. If these particles remain about the asteroid, they could potentially jeopardize the mission as, in the event of a collision with the Hayabusa2 spacecraft, the spacecraft’s functionality could be reduced. In this paper, we make use of invariant manifold theory to assess the conditions - impact location, particle radius, ejection velocity - that cause ejecta particles to get captured in periodic orbits. The analysis is carried out within the dynamical framework of the perturbed Augmented Hill Problem, which takes into account the solar radiation pressure, the effect of eclipses, and the J2 and J4 terms of the asteroid’s gravity potential in its spherical harmonics expansion. We analyze millimeter to centimeter sized particles and captures into three families of periodic orbits that are robust to large values of the solar radiation pressure acceleration – the traditional a and g’ families of the Hill Problem and the southern halo orbits. We go on to find the impact locations for the SCI from where ejecta particles are most likely to be captured into periodic orbits via their stable manifolds. As such, we recover the sets of initial states that lead ejecta to temporary orbital capture and show that solar radiation pressure cannot be neglected in these analyses, identifying locations on the Sun side of the asteroid at medium latitudes as the best impact locations.
Advisors/Committee Members: Heiligers, Jeannette (mentor), Soldini, Stefania (mentor), Delft University of Technology (degree granting institution).
Subjects/Keywords: Augmented Hill Problem; Astrodynamics; Asteroid Ejecta
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APA (6th Edition):
Villegas Pinto, D. (. (2019). Temporary Capture of Asteroid Ejecta into Periodic Orbits: Application to JAXA’s Hayabusa2 Impact Event. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:d9768a7f-0c57-4437-befb-4ccf7a39a0a5
Chicago Manual of Style (16th Edition):
Villegas Pinto, Daniel (author). “Temporary Capture of Asteroid Ejecta into Periodic Orbits: Application to JAXA’s Hayabusa2 Impact Event.” 2019. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:d9768a7f-0c57-4437-befb-4ccf7a39a0a5.
MLA Handbook (7th Edition):
Villegas Pinto, Daniel (author). “Temporary Capture of Asteroid Ejecta into Periodic Orbits: Application to JAXA’s Hayabusa2 Impact Event.” 2019. Web. 04 Mar 2021.
Vancouver:
Villegas Pinto D(. Temporary Capture of Asteroid Ejecta into Periodic Orbits: Application to JAXA’s Hayabusa2 Impact Event. [Internet] [Masters thesis]. Delft University of Technology; 2019. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:d9768a7f-0c57-4437-befb-4ccf7a39a0a5.
Council of Science Editors:
Villegas Pinto D(. Temporary Capture of Asteroid Ejecta into Periodic Orbits: Application to JAXA’s Hayabusa2 Impact Event. [Masters Thesis]. Delft University of Technology; 2019. Available from: http://resolver.tudelft.nl/uuid:d9768a7f-0c57-4437-befb-4ccf7a39a0a5
Delft University of Technology
17.
Hess, J.R. (author).
Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists.
Degree: 2016, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:dee2d842-ca91-4086-8ef8-771cfee1b0dc
► Interplanetary missions use gravitational slingshots around planetary bodies to adjust their heliocentric velocity or inclination for quite some time. The momentum exchange that can be…
(more)
▼ Interplanetary missions use gravitational slingshots around planetary bodies to adjust their heliocentric velocity or inclination for quite some time. The momentum exchange that can be achieved during a so-called gravity assist is limited by the mass of the planetary body. To overcome this limitation, an aerogravity assist was proposed, a maneuver where, in addition to the gravitational forces, use is made of aerodynamic forces to increase the bending angle of the velocity, hence increasing the momentum exchange. To investigate how efficient an aerogravity assist can change the interplanetary orbital inclination and velocity, a simulator was developed that is capable of simulating both the gravitational and aerodynamic forces on a vehicle during an aerogravity assist. It was determined that waverider is a type of vehicle suitable for aerogravity assists due to their large lift-to-drag ratio, which reduces the energy dissipation in the atmosphere. The aerodynamic characteristics of a number of waverider shapes were evaluated, after which the one with the largest lift-to-drag ratio was selected. Furthermore, a numerical optimization algorithm was used to develop a reference trajectory planner. Finally, a guidance algorithm based on the tracking of drag accelerations was developed and tested to investigate if the found trajectories would still be feasible under the influence of uncertainties and perturbations. The angle over which the trajectory is bent is a measure for the effectiveness of the aerogravity assist. Using the reference trajectory planner, the maximum possible atmospheric bending angle was investigated for an aerogravity assist at Mars and Jupiter for different initial velocities. From this analysis, it was concluded that extremely high velocities were involved in the aerogravity assist at Jupiter, which resulted in large mechanical and thermal loads. These loads would limit the achievable bending angle when the velocities become too large. For the entry velocities investigated, the velocity bending angle could be increased by 10% for high entry (80.0 km/s) velocities and up to 143% for a relatively low entry velocity (68.0 km/s). For an entry velocity of 80.0 km/s, the initial heat-flux peak exceeded the imposed constraints, which prevented the optimization algorithm of finding any solutions. The maximum velocity bending angle that could be achieved at Jupiter was 125.1 degrees at an entry velocity of 68.0 km/s. At Mars, although the heat loads were still larger than for an Earth entry, it is believed that thermal protection systems can be designed that could handle the heat loads. The velocity bending angle could be increased by 490% to 818% depending on the arrival velocity, with a maximum velocity bending angle of 178.5 degrees at an entry velocity of 9.0 km/s. To investigate the effect of an aerogravity assist on an actual mission, two existing missions has been selected: Rosetta for Mars and Ulysses for Jupiter. Although both spacecraft did not have an aerodynamic shape, which means an aerogravity…
Advisors/Committee Members: Mooij, E. (mentor), Sudmeijer, K.J. (mentor).
Subjects/Keywords: astrodynamics; aerogravity; gravity; assist; optimization; hypersonics; waverider
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APA (6th Edition):
Hess, J. R. (. (2016). Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:dee2d842-ca91-4086-8ef8-771cfee1b0dc
Chicago Manual of Style (16th Edition):
Hess, J R (author). “Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists.” 2016. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:dee2d842-ca91-4086-8ef8-771cfee1b0dc.
MLA Handbook (7th Edition):
Hess, J R (author). “Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists.” 2016. Web. 04 Mar 2021.
Vancouver:
Hess JR(. Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists. [Internet] [Masters thesis]. Delft University of Technology; 2016. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:dee2d842-ca91-4086-8ef8-771cfee1b0dc.
Council of Science Editors:
Hess JR(. Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists. [Masters Thesis]. Delft University of Technology; 2016. Available from: http://resolver.tudelft.nl/uuid:dee2d842-ca91-4086-8ef8-771cfee1b0dc
Cal Poly
18.
Iuliano, Jay R.
A Solution to the Circular Restricted N Body Problem in Planetary Systems.
Degree: MS, Aerospace Engineering, 2016, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/1612
;
10.15368/theses.2016.80
► This thesis is a brief look at a new solution to a problem that has been approached in many different ways in the past…
(more)
▼ This thesis is a brief look at a new solution to a problem that has been approached in many different ways in the past - the N body problem. By focusing on planetary systems, satellite dynamics can be modeled in a fashion similar to the Circular Restricted Three Body Problem (CR3BP) with the Circular Restricted N Body Problem (CRNBP). It was found that this new formulation of the dynamics can then utilize the tools created from all the research into the CR3BP to reassess the possibility of different complex trajectories in systems where there are more than just two large gravitational bodies affecting the dynamics, namely periodic and semi-periodic orbits, halo orbits, and low energy transfers It was also found that not only system dynamics, but models of the Jacobi constant could also be formulated similarly to the CR3BP. Validating the authenticity of these new sets of equations, the CRNBP dynamics are applied to a satellite in the Earth-Moon system and compared to a simulation of the CR3BP under identical circumstances. This test verified the dynamics of the CRNBP, showing that the two systems created almost identical results with relatively small deviations over time and with essentially identical path trends. In the Jovian system, it was found the mass ratio required to validated the assumptions required to integrate the equations of motion was around .1%. Once the mass ratio grew past that limit, trajectories propagated with the CRNBP showed significant deviation from trajectories propagated with a higher fidelity model of Newtonian motion. The results from the derivation of the Jacobi constant are consistent with the 3 body system, but they are fairly standalone.
Advisors/Committee Members: Kira Abercromby.
Subjects/Keywords: Astrodynamics; Navigation, Guidance, Control and Dynamics
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APA (6th Edition):
Iuliano, J. R. (2016). A Solution to the Circular Restricted N Body Problem in Planetary Systems. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/1612 ; 10.15368/theses.2016.80
Chicago Manual of Style (16th Edition):
Iuliano, Jay R. “A Solution to the Circular Restricted N Body Problem in Planetary Systems.” 2016. Masters Thesis, Cal Poly. Accessed March 04, 2021.
https://digitalcommons.calpoly.edu/theses/1612 ; 10.15368/theses.2016.80.
MLA Handbook (7th Edition):
Iuliano, Jay R. “A Solution to the Circular Restricted N Body Problem in Planetary Systems.” 2016. Web. 04 Mar 2021.
Vancouver:
Iuliano JR. A Solution to the Circular Restricted N Body Problem in Planetary Systems. [Internet] [Masters thesis]. Cal Poly; 2016. [cited 2021 Mar 04].
Available from: https://digitalcommons.calpoly.edu/theses/1612 ; 10.15368/theses.2016.80.
Council of Science Editors:
Iuliano JR. A Solution to the Circular Restricted N Body Problem in Planetary Systems. [Masters Thesis]. Cal Poly; 2016. Available from: https://digitalcommons.calpoly.edu/theses/1612 ; 10.15368/theses.2016.80
Cal Poly
19.
Jackson, Daniel J.
Formulation of an Optimal Search Strategy for Space Debris At GEO.
Degree: MS, Aerospace Engineering, 2011, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/656
;
10.15368/theses.2011.202
► The purpose of this thesis is to create a search strategy to find orbital debris when the object fails to appear in the sky…
(more)
▼ The purpose of this thesis is to create a search strategy to find orbital debris when the object fails to appear in the sky at its predicted location. This project is for NASA Johnson Space Center Orbital Debris Program Office through the MODEST (Michigan Orbital Debris Survey Telescope) program. This thesis will build upon the research already done by James Biehl in “Formulation of a Search Strategy for Space Debris at GEO.” MODEST tracks objects at a specific right ascension and declination. A circular orbit assumption is then used to predict the location of the object at a later time. Another telescope performs a follow-up to the original observation to provide a more accurate orbit predication. This thesis develops a search strategy when the follow-up is not successful. A general search strategy for finding space debris was developed based on previous observations. A GUI was also generated to find a search strategy in real-time for a specific object based upon previous observations of that object.
Search strategies were found by adding a 2% mean random error to the position and velocity vectors. Adding a random error allows for finding the most likely location of space debris when the orbital elements are slightly incorrect. A bivariate kernel density estimator was used to find the probability density function. The probability density function was used to find the most probable location of an object. A correlation between error in the orbital elements and error in right ascension and declination root mean square (RMS) error was investigated. It was found that the orbital elements affect the RMS error nonlinearly, but the relation between orbital element and error depended on the object and no general pattern was found. It was found that how long after the original object was found until the follow-up was attempted did not have a large impact on the probability density function or the search strategy.
Advisors/Committee Members: Kira Abercromby.
Subjects/Keywords: aerospace; space; debris; GEO; telescope; MODEST; Astrodynamics
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APA (6th Edition):
Jackson, D. J. (2011). Formulation of an Optimal Search Strategy for Space Debris At GEO. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/656 ; 10.15368/theses.2011.202
Chicago Manual of Style (16th Edition):
Jackson, Daniel J. “Formulation of an Optimal Search Strategy for Space Debris At GEO.” 2011. Masters Thesis, Cal Poly. Accessed March 04, 2021.
https://digitalcommons.calpoly.edu/theses/656 ; 10.15368/theses.2011.202.
MLA Handbook (7th Edition):
Jackson, Daniel J. “Formulation of an Optimal Search Strategy for Space Debris At GEO.” 2011. Web. 04 Mar 2021.
Vancouver:
Jackson DJ. Formulation of an Optimal Search Strategy for Space Debris At GEO. [Internet] [Masters thesis]. Cal Poly; 2011. [cited 2021 Mar 04].
Available from: https://digitalcommons.calpoly.edu/theses/656 ; 10.15368/theses.2011.202.
Council of Science Editors:
Jackson DJ. Formulation of an Optimal Search Strategy for Space Debris At GEO. [Masters Thesis]. Cal Poly; 2011. Available from: https://digitalcommons.calpoly.edu/theses/656 ; 10.15368/theses.2011.202
University of New South Wales
20.
Capon, Christopher.
Ionospheric Aerodynamics in Low Earth Orbit.
Degree: Engineering & Information Technology, 2017, University of New South Wales
URL: http://handle.unsw.edu.au/1959.4/58652
;
https://unsworks.unsw.edu.au/fapi/datastream/unsworks:46528/SOURCE01?view=true
► Understanding perturbing forces takes on new significance as the Low Earth Orbit (LEO) environment becomes increasingly congested and the risk of collision events that threaten…
(more)
▼ Understanding perturbing forces takes on new significance as the Low Earth Orbit (LEO) environment becomes increasingly congested and the risk of collision events that threaten access to space infrastructure grows. The perturbation caused by the charged aerodynamic interaction of LEO objects with the ionosphere ("ionospheric aerodynamics'') is currently poorly understood and not modelled accurately. The work presented in this thesis provides quantitative insight into the physics underpinning ionospheric aerodynamics and its influence on the orbital motion of LEO objects.By deriving the set of scaling parameters that describe the electrostatic interaction of a K-species plasma with a charged object, this work has reduced the 4+5K quantities that define plasma interactions to 1+4K independent scaling parameters, and in doing so has helped make the study of charged ionospheric aerodynamics feasible. These scaling parameters represent a significant generalisation of previous work, including linking high and low surface potential plasma phenomena through a new general plasma shielding length lambda_phi. Plasma interaction phenomena that influence charged aerodynamics are then represented in a two-dimensional phase-space "P" defined by two key dimensionless scaling parameters: the ion deflection parameter "alpha" and the general shielding ratio "chi". A map of plasma interaction phenomena within "P" was developed and related to charged aerodynamic forces providing new insights into phenomena that govern direct and indirect charged aerodynamic forces. This map was then applied to develop a physics-based framework to predict the influence of ionospheric aerodynamics on LEO objects.This work predicts that ionospheric aerodynamic forces may represent up to 0.05-18% of the total aerodynamic force vector experienced by a cylindrical object with a surface potential between -0.75 V and -30 V at 500 km altitude, increasing to 0.9-78% at 1500 km. Therefore, this work concludes that ionospheric aerodynamics can have a significant influence on the motion of LEO objects.
Advisors/Committee Members: Boyce, Russell, Engineering & Information Technology, UNSW Canberra, UNSW, Brown, Melrose, Engineering & Information Technology, UNSW Canberra, UNSW.
Subjects/Keywords: Astrodynamics; Ionospheric Aerodynamics; Space Situational Awareness
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APA (6th Edition):
Capon, C. (2017). Ionospheric Aerodynamics in Low Earth Orbit. (Doctoral Dissertation). University of New South Wales. Retrieved from http://handle.unsw.edu.au/1959.4/58652 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:46528/SOURCE01?view=true
Chicago Manual of Style (16th Edition):
Capon, Christopher. “Ionospheric Aerodynamics in Low Earth Orbit.” 2017. Doctoral Dissertation, University of New South Wales. Accessed March 04, 2021.
http://handle.unsw.edu.au/1959.4/58652 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:46528/SOURCE01?view=true.
MLA Handbook (7th Edition):
Capon, Christopher. “Ionospheric Aerodynamics in Low Earth Orbit.” 2017. Web. 04 Mar 2021.
Vancouver:
Capon C. Ionospheric Aerodynamics in Low Earth Orbit. [Internet] [Doctoral dissertation]. University of New South Wales; 2017. [cited 2021 Mar 04].
Available from: http://handle.unsw.edu.au/1959.4/58652 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:46528/SOURCE01?view=true.
Council of Science Editors:
Capon C. Ionospheric Aerodynamics in Low Earth Orbit. [Doctoral Dissertation]. University of New South Wales; 2017. Available from: http://handle.unsw.edu.au/1959.4/58652 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:46528/SOURCE01?view=true
Delft University of Technology
21.
Carvalho Barreira Lages Ribeiro, André (author).
Optimal Aeroassisted Maneuvers for Orbital Transfer.
Degree: 2019, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398
► Orbital plane changes require a considerable amount of propellant to be completed if the traditional fully propulsive maneuver is used. The cost associated with this…
(more)
▼ Orbital plane changes require a considerable amount of propellant to be completed if the traditional fully propulsive maneuver is used. The cost associated with this type of maneuvers is one of its major drawbacks, and performing an inclination change in orbit would require a large part of the total mass budget of the mission. Moreover, launching such vehicles or satellites would demand the usage of a larger launcher, further contributing to the increase of the mission cost. An alternative to the fully propulsive maneuver is the aeroassisted maneuver, proposed London (1961), where the vehicle performs one or more passes through the atmosphere of the central body. This allows for the aerodynamic forces to steer and brake the vehicle, reducing its orbital speed and changing the inclination of the orbit with less propellant required. To investigate the impact of the aeroassisted maneuver in the propellant requirements, the following research question was proposed: What is the impact of using an aeroassisted maneuver in reducing the amount of propellant needed to achieve a certain orbital inclination change? A simulation environment was developed using the TU Delft
Astrodynamics Toolbox, capable of simulating an aeroassisted trajectory from a Geostationary Earth Orbit to a Low Earth Orbit, with an associated inclination change. A simple lateral guidance algorithm, capable of tracking the orbital plane, was also developed, and a node control algorithm was implemented to chose the guidance nodes during each individual trajectory. The performance of three vehicle configurations, with different lift-to-drag ratios, were investigated in terms of achievable inclination change. The HORUS-2B, a vehicle with a moderate to high lift-to-drag ratio, was the most suitable to perform the maneuver, as it allows for a larger inclination change while reducing the energy dissipation rate in the atmosphere. The simulator was integrated with an optimization algorithm, such that optimal trajectories could be investigated. The performance of the aeroassisted maneuver is measured in terms of three different objectives: the offset in inclination between the final and the target orbital planes, the velocity impulse applied at atmospheric exit to target the desired apogee and the maximum heat load experienced by the vehicle during any given atmospheric pass. The NSGA-II optimizer is selected to minimize the three objectives, and several parameters are tuned to increase the probability of finding the global optimum, namely the tuning parameters of the optimizer itself, the values of the constraint and the number of guidance nodes. It was found that an orbital plane of 20 degrees could be targeted with an error of only 0.28%, with the total maneuver requiring 87.32% less Delta V needed from thrust impulses. The heat load obtained is 1315.6 kJ/m2, which is well within the allowable range for Thermal Protection Systems currently available, and the total number of atmospheric passes is five. Moreover, it is possible to…
Advisors/Committee Members: Mooij, Erwin (mentor), Delft University of Technology (degree granting institution).
Subjects/Keywords: Optimization; Re-entry; Aeroassisted; Astrodynamics; Hypersonic
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APA (6th Edition):
Carvalho Barreira Lages Ribeiro, A. (. (2019). Optimal Aeroassisted Maneuvers for Orbital Transfer. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398
Chicago Manual of Style (16th Edition):
Carvalho Barreira Lages Ribeiro, André (author). “Optimal Aeroassisted Maneuvers for Orbital Transfer.” 2019. Masters Thesis, Delft University of Technology. Accessed March 04, 2021.
http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398.
MLA Handbook (7th Edition):
Carvalho Barreira Lages Ribeiro, André (author). “Optimal Aeroassisted Maneuvers for Orbital Transfer.” 2019. Web. 04 Mar 2021.
Vancouver:
Carvalho Barreira Lages Ribeiro A(. Optimal Aeroassisted Maneuvers for Orbital Transfer. [Internet] [Masters thesis]. Delft University of Technology; 2019. [cited 2021 Mar 04].
Available from: http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398.
Council of Science Editors:
Carvalho Barreira Lages Ribeiro A(. Optimal Aeroassisted Maneuvers for Orbital Transfer. [Masters Thesis]. Delft University of Technology; 2019. Available from: http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398
University of Colorado
22.
McMahon, Jay Warren.
An Analytical Theory for the Perturbative Effect of Solar Radiation Pressure on Natural and Artificial Satellites.
Degree: PhD, Aerospace Engineering Sciences, 2011, University of Colorado
URL: https://scholar.colorado.edu/asen_gradetds/26
► Solar radiation pressure is the largest non-gravitational perturbation for most satellites in the solar system, and can therefore have a significant influence on their…
(more)
▼ Solar radiation pressure is the largest non-gravitational perturbation for most satellites in the solar system, and can therefore have a significant influence on their orbital dynamics. This work presents a new method for representing the solar radiation pressure force acting on a satellite, and applies this theory to natural and artificial satellites. The solar radiation pressure acceleration is modeled as a Fourier series which depends on the Sun's location in a body-fixed frame; a new set of Fourier coefficients are derived for every latitude of the Sun in this frame, and the series is expanded in terms of the longitude of the Sun. The secular effects due to the solar radiation pressure perturbations are given analytically through the application of averaging theory when the satellite is in a synchronous orbit. This theory is then applied to binary asteroid systems to explain the Binary YORP effect. Long term predictions of the evolution of the near-Earth asteroid 1999 KW4 are discussed under the influence of solar radiation pressure, J2, and 3rd body gravitational effects from the Sun. Secular effects are shown to remain when the secondary asteroid becomes non-synchronous due to a librational motion. The theory is also applied to Earth orbiting spacecraft, and is shown to be a valuable tool for improved orbit determination. The Fourier series solar radiation pressure model derived here is shown to give comparable results for orbit determination of the GPS IIR-M satellites as JPL's solar radiation pressure model. The theory is also extended to incorporate the effects of the Earth's shadow analytically. This theory is briefly applied to the evolution of orbital debris to explain the assumptions that are necessary in order to use the cannonball model for debris orbit evolution, as is common in the literature. Finally, the averaging theory methodology is applied to a class of Earth orbiting solar sail spacecraft to show the orbital effects when the sails are made to cone at orbit rates in the local horizontal frame.
Advisors/Committee Members: Daniel J. Scheeres, Hanspeter Schaub, George Born.
Subjects/Keywords: astrodynamics; BYORP; celestial mechanics; orbital dynamics; orbit determination; solar radiation pressure; Aerospace Engineering; Astrodynamics
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APA (6th Edition):
McMahon, J. W. (2011). An Analytical Theory for the Perturbative Effect of Solar Radiation Pressure on Natural and Artificial Satellites. (Doctoral Dissertation). University of Colorado. Retrieved from https://scholar.colorado.edu/asen_gradetds/26
Chicago Manual of Style (16th Edition):
McMahon, Jay Warren. “An Analytical Theory for the Perturbative Effect of Solar Radiation Pressure on Natural and Artificial Satellites.” 2011. Doctoral Dissertation, University of Colorado. Accessed March 04, 2021.
https://scholar.colorado.edu/asen_gradetds/26.
MLA Handbook (7th Edition):
McMahon, Jay Warren. “An Analytical Theory for the Perturbative Effect of Solar Radiation Pressure on Natural and Artificial Satellites.” 2011. Web. 04 Mar 2021.
Vancouver:
McMahon JW. An Analytical Theory for the Perturbative Effect of Solar Radiation Pressure on Natural and Artificial Satellites. [Internet] [Doctoral dissertation]. University of Colorado; 2011. [cited 2021 Mar 04].
Available from: https://scholar.colorado.edu/asen_gradetds/26.
Council of Science Editors:
McMahon JW. An Analytical Theory for the Perturbative Effect of Solar Radiation Pressure on Natural and Artificial Satellites. [Doctoral Dissertation]. University of Colorado; 2011. Available from: https://scholar.colorado.edu/asen_gradetds/26
University of Colorado
23.
Baresi, Nicola.
Spacecraft Formation Flight on Quasi-Periodic Invariant Tori.
Degree: PhD, 2017, University of Colorado
URL: https://scholar.colorado.edu/asen_gradetds/219
► Since the successful rendezvous of the Gemini VI and VII spacecraft in 1965, spacecraft formation flying has attracted the interest of many researchers in…
(more)
▼ Since the successful rendezvous of the Gemini VI and VII spacecraft in 1965, spacecraft formation flying has attracted the interest of many researchers in the field. Yet, existing methodologies do not currently account for the oblateness of a central body when the distance between the satellites exceeds the reach of standard analytical techniques such as Brouwer-Lyddane theory and thereof. In this dissertation, the problem of designing bounded relative orbits is approached with a dynamical systems theory perspective in order to overcome the limitations imposed by mean-to-osculating orbit element mappings and linearization errors. We find that the dynamics of satellites in the Earth zonal problem can be fundamentally described by three periods, whose averaged values can be accurately computed through numerical integration. To ensure long-term bounded relative motion between the satellites in a formation, at least two of their fundamental periods need to be matched on average. This condition is enforced by including additional constraints into existing techniques for calculating families of quasi-periodic invariant tori. The result is a numerical procedure that searches for the invariant curves of a stroboscopic mapping while changing the polar component of the angular momentum vector for each of the quasi-periodic tori within the family. Upon convergence, the algorithm outputs several curves that can be interpolated to obtain an entire surface of bounded relative motion. That is, by selecting arbitrary initial conditions on this surface, bounded relative motion can be established, regardless of the number of zonal harmonics terms that are included in the geopotential. Given this encouraging result, we move beyond Earth's orbit and investigate the problem of designing bounded relative orbits about small irregular bodies. First, we consider the case of asteroid (4179) Toutatis, and build on previous research to identify periodic and quasi-periodic orbits that ensure boundedness in spite of the complex shape and rotational state of the target asteroid. Next, we move to the Martian system and design spacecraft formations near Phobos. Once again, we aim at improving the realism of previous simulations found in the literature by modeling the nonspherical shape and nonzero eccentricity of the Martian moon. The resulting higher fidelity model causes entire families of periodic orbits to become quasi-periodic invariant tori that eventually serve as initial conditions for bounded spacecraft formations. The last part of this thesis is dedicated to assessing the robustness of the relative trajectories computed throughout the manuscript. Although atmospheric drag and solar radiation pressure have catastrophic effects on the relative dynamics of satellites in LEO and near Toutatis, it is found that spacecraft formations in MEO, GEO, and about Phobos are quite resilient to mismodeled dynamics, making quasi-periodic invariant tori a robust option for flying satellite clusters in these complex dynamical…
Advisors/Committee Members: Daniel J. Scheeres, Natasha Bosanac, Jay W. McMahon, James D. Miss, Hanspeter Schaub.
Subjects/Keywords: astrodynamics; cluster flight; dynamical systems theory; quasi-periodic invariant tori; spacecraft formation flying; Aerospace Engineering; Astrodynamics
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APA (6th Edition):
Baresi, N. (2017). Spacecraft Formation Flight on Quasi-Periodic Invariant Tori. (Doctoral Dissertation). University of Colorado. Retrieved from https://scholar.colorado.edu/asen_gradetds/219
Chicago Manual of Style (16th Edition):
Baresi, Nicola. “Spacecraft Formation Flight on Quasi-Periodic Invariant Tori.” 2017. Doctoral Dissertation, University of Colorado. Accessed March 04, 2021.
https://scholar.colorado.edu/asen_gradetds/219.
MLA Handbook (7th Edition):
Baresi, Nicola. “Spacecraft Formation Flight on Quasi-Periodic Invariant Tori.” 2017. Web. 04 Mar 2021.
Vancouver:
Baresi N. Spacecraft Formation Flight on Quasi-Periodic Invariant Tori. [Internet] [Doctoral dissertation]. University of Colorado; 2017. [cited 2021 Mar 04].
Available from: https://scholar.colorado.edu/asen_gradetds/219.
Council of Science Editors:
Baresi N. Spacecraft Formation Flight on Quasi-Periodic Invariant Tori. [Doctoral Dissertation]. University of Colorado; 2017. Available from: https://scholar.colorado.edu/asen_gradetds/219
University of Colorado
24.
De Smet, Stijn.
On the Design of Solar Gravity Driven Planetocentric Transfers Using Artificial Neural Networks.
Degree: PhD, 2018, University of Colorado
URL: https://scholar.colorado.edu/asen_gradetds/235
► The sun's gravity can be used to efficiently transfer between different planetocentric orbits. Such transfers cannot be designed in a two-body dynamical system, nor…
(more)
▼ The sun's gravity can be used to efficiently transfer between different planetocentric orbits. Such transfers cannot be designed in a two-body dynamical system, nor do analytical methods exist to identify such transfers. This dissertation presents a method to efficiently identify transfers between a specified departure and target orbit. This method is applied to a well known problem: transfers from inclined low-earth orbits to the geostationary orbit. Motivated by the large observed control authority of the sun for geocentric transfers, a new mission architecture is defined. This architecture allows the injection of multiple spacecraft around Mars in different target orbits, enabled by solar gravity driven orbital transfers. The efficient design of applications for a wide variety of departure and target orbits, requires an understanding of a large area of the phase space. This dissertation showcases how an artificial neural network architecture can accurately predict the solar gravity driven transfers, for a significantly large section of the phase space. The developed architecture is then used to efficiently identify transfers for several different applications. Multiple revolution transfers with maneuvers at intermediate periareions are identified that arrive at Phobos or Deimos. Furthermore, transfers are designed that transfer to both Phobos and Deimos in a single trajectory. In addition to addressing solar perturbed planetocentric transfers, this dissertation shows how the developed artificial neural network framework can be applied to a different problem, with different dynamics. As an example, the dissertation develops an artificial neural network architecture that can predict heteroclinic connections in the Earth-Moon circular restricted three-body problem.
Advisors/Committee Members: Daniel J. Scheeres, Natasha Bosanac, Jay McMahon, James D. Meiss, Jeffrey S. Parker.
Subjects/Keywords: artificial neural networks; astrodynamics; eccentric hill system; machine learning; periapse poincaré maps; Aerospace Engineering; Astrodynamics; Computer Sciences
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APA (6th Edition):
De Smet, S. (2018). On the Design of Solar Gravity Driven Planetocentric Transfers Using Artificial Neural Networks. (Doctoral Dissertation). University of Colorado. Retrieved from https://scholar.colorado.edu/asen_gradetds/235
Chicago Manual of Style (16th Edition):
De Smet, Stijn. “On the Design of Solar Gravity Driven Planetocentric Transfers Using Artificial Neural Networks.” 2018. Doctoral Dissertation, University of Colorado. Accessed March 04, 2021.
https://scholar.colorado.edu/asen_gradetds/235.
MLA Handbook (7th Edition):
De Smet, Stijn. “On the Design of Solar Gravity Driven Planetocentric Transfers Using Artificial Neural Networks.” 2018. Web. 04 Mar 2021.
Vancouver:
De Smet S. On the Design of Solar Gravity Driven Planetocentric Transfers Using Artificial Neural Networks. [Internet] [Doctoral dissertation]. University of Colorado; 2018. [cited 2021 Mar 04].
Available from: https://scholar.colorado.edu/asen_gradetds/235.
Council of Science Editors:
De Smet S. On the Design of Solar Gravity Driven Planetocentric Transfers Using Artificial Neural Networks. [Doctoral Dissertation]. University of Colorado; 2018. Available from: https://scholar.colorado.edu/asen_gradetds/235
Texas A&M University
25.
Johnson, Kirk Wayne.
Approaches for Modeling Satellite Relative Motion.
Degree: PhD, Aerospace Engineering, 2016, Texas A&M University
URL: http://hdl.handle.net/1969.1/158945
► This dissertation explores new approaches for modeling perturbed and unperturbed satellite relative motion. It extends Hoots orbit theory, an analytical averaging-method perturbation solution to the…
(more)
▼ This dissertation explores new approaches for modeling perturbed and unperturbed satellite relative motion. It extends Hoots orbit theory, an analytical averaging-method perturbation solution to the Zonal Problem, to second order. In addition, this study develops a new hybrid numerical/analytical algorithm for converting initial conditions from osculating elements to mean elements, so that a single set of osculating initial conditions may be taken as simulation inputs. Also, this study develops a new version of the Gim-Alfriend State Transition Matrix (GA STM) for linearized perturbed relative motion, in terms of the variables from Hoots theory. These variables, the Hoots elements, are advantageous (although not unique) in that they have no singularities for orbit eccentricity or inclination and they require only one solution of Kepler's Equation at each time step, even when using the GA STM. The new models are compared by simulation with orbit theories and GA STMs using the so-called nonsingular elements (which are in fact singular for zero inclination) and the equinoctial elements. This study predicts and verifies the order of magnitude of modeling error due to various sources.
This study also considers two special applications in satellite relative motion modeling. First, Projected Circular Orbit (PCO) formations, originally defined for unperturbed motion about a circular reference orbit, have important applications and are widely studied. This dissertation removes the singularity for zero inclination by implementing the PCO initial conditions in equinoctial elements, allowing PCO formations to be initialized about equatorial orbits. Furthermore, this study reveals how the choice of variables for writing the PCO initial conditions changes the geometric interpretation of the PCO phase angle parameter α. Second, this study develops an alternative to the standard methods for mitigating along-track drift in perturbed satellite formations. The new method eliminates all along-track secular motion to first order by sacrificing one degree of freedom in the formation design.
Advisors/Committee Members: Alfriend, Kyle T. (advisor), Vadali, Srinivas R. (advisor), Hurtado, John E. (committee member), Zelenko, Igor (committee member).
Subjects/Keywords: Satellite Relative Motion; Satellite Formation Flying; Zonal Problem; Astrodynamics
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APA (6th Edition):
Johnson, K. W. (2016). Approaches for Modeling Satellite Relative Motion. (Doctoral Dissertation). Texas A&M University. Retrieved from http://hdl.handle.net/1969.1/158945
Chicago Manual of Style (16th Edition):
Johnson, Kirk Wayne. “Approaches for Modeling Satellite Relative Motion.” 2016. Doctoral Dissertation, Texas A&M University. Accessed March 04, 2021.
http://hdl.handle.net/1969.1/158945.
MLA Handbook (7th Edition):
Johnson, Kirk Wayne. “Approaches for Modeling Satellite Relative Motion.” 2016. Web. 04 Mar 2021.
Vancouver:
Johnson KW. Approaches for Modeling Satellite Relative Motion. [Internet] [Doctoral dissertation]. Texas A&M University; 2016. [cited 2021 Mar 04].
Available from: http://hdl.handle.net/1969.1/158945.
Council of Science Editors:
Johnson KW. Approaches for Modeling Satellite Relative Motion. [Doctoral Dissertation]. Texas A&M University; 2016. Available from: http://hdl.handle.net/1969.1/158945
Penn State University
26.
Paik, Ghanghoon.
Optimal Orbit Raising Via Particle Swarm Optimization
.
Degree: 2015, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/25366
► A spacecraft in one orbit may need to move to another orbit. Apoapsis orbit raising, in particular, takes a spacecraft from a circular orbit to…
(more)
▼ A spacecraft in one orbit may need to move to another orbit. Apoapsis orbit raising, in particular, takes a spacecraft from a circular orbit to an elliptical orbit by thrusting at a periapsis. This technique was applied as the initial stage for the lunar (LADEE), GEO (ARTEMIS), and interplanetary (Mangalyaan) missions to save propellant usage and raise the apoapsis distance of the orbits. These projects show that the apoapsis orbit raising can be applied to various types of missions.
In this thesis, by applying the Particle Swarm Optimization (PSO) algorithm to the five finite thrust maneuvers, evaluation of an optimal solution that derives optimized propellant usage is presented. Each transfer orbit pushes out the apoapsis of the trajectory, depending on the thrust duration and the thrust-on location. The final orbit of the optimal solution of the problem should meet two criteria: the line of apside (LOA) alignment and the apoapsis distance.
The PSO is a computational method that is inspired by a swarm movement. By sharing information obtained by each member, the entire swarm (set of possible solutions) can find the best location efficiently and rapidly. This algorithm highly depends on size of a swarm and number of iterations as well as an initial solution set. In this thesis, a modification is applied to the PSO to handle equality constraints.
The PSO application to apoapsis orbit raising shows the feasibility of determining an optimized trajectory to reach a target orbit and gives required propellant for each maneuver in terms of a thrust duration. The optimal results are acquired by the PSO algorithm and all the requirements are satisfied.
Advisors/Committee Members: Robert Graham Melton, Thesis Advisor/Co-Advisor.
Subjects/Keywords: Astrodynamics; Orbit Raising; Particle Swarm Optimization; Trajectory Optimization
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APA ·
Chicago ·
MLA ·
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CSE |
Export
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APA (6th Edition):
Paik, G. (2015). Optimal Orbit Raising Via Particle Swarm Optimization
. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/25366
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Paik, Ghanghoon. “Optimal Orbit Raising Via Particle Swarm Optimization
.” 2015. Thesis, Penn State University. Accessed March 04, 2021.
https://submit-etda.libraries.psu.edu/catalog/25366.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Paik, Ghanghoon. “Optimal Orbit Raising Via Particle Swarm Optimization
.” 2015. Web. 04 Mar 2021.
Vancouver:
Paik G. Optimal Orbit Raising Via Particle Swarm Optimization
. [Internet] [Thesis]. Penn State University; 2015. [cited 2021 Mar 04].
Available from: https://submit-etda.libraries.psu.edu/catalog/25366.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Paik G. Optimal Orbit Raising Via Particle Swarm Optimization
. [Thesis]. Penn State University; 2015. Available from: https://submit-etda.libraries.psu.edu/catalog/25366
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Penn State University
27.
Conte, Davide.
Semi-Analytical Solutions for Proximity Operations in the Circular Restricted Three-Body Problem.
Degree: 2019, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/16596dxc5120
► The research presented in this dissertation aims at characterizing the relative motion between spacecraft in periodic orbits in the circular restricted three-body problem. Proximity operations…
(more)
▼ The research presented in this dissertation aims at characterizing the relative motion between spacecraft in periodic orbits in the circular restricted three-body problem. Proximity operations maneuvers, such as rendezvous and station-keeping, are approximated using a semi-analytical approach, i.e. by combining the analytical approximation of the nominal periodic orbit of the targeted spacecraft or targeted orbital location with simplified equations of motion that are derived by assuming that the chasing spacecraft and target (or targeted location) are always "close" to each other. The presented method is compared to the "exact" solutions obtained by numerically integrating the non-linear equations of motion of the circular restricted three-body problem. Orbit propagation examples, i.e. control-free trajectories, are shown along with proximity operation maneuvers for which delta-v's are computed. Suitable initial conditions for proximity operations are also computed based on known orbital transfers to specific orbits of interest, including halo orbits and distant retrograde orbits in the Earth-Moon and Mars-Phobos systems. Propellant-optimal results are found and compared to nearby local minima for a given set of time constraints in order to find a maneuver that allows for flexibility regarding departure and arrival time along with contingency plans. This approximation is validated against the use of the more computationally expensive full nonlinear equations of motion of the three-body problem and the "area of applicability" of this method is defined based on a metric that takes into account the initial conditions used when initiating proximity operations and the time-of-flight required to accomplish such maneuvers. Sample results for the Earth-Moon and Mars-Phobos systems are presented for cis-lunar and cis-Martian orbits of interest. The implementation of this method in pre-phase A mission design is also demonstrated in a sample end-to-end Earth-to-Mars mission. The method presented in this dissertation is shown to accurately describe the control-free relative motion between spacecraft in addition to being able to predict the necessary delta-v maneuvers for various proximity operations. Additionally, this method requires less computational time than full numerical methods while being able to assess its accuracy and the validity of the results obtained. Limitations of this method are imposed on the initial relative position between spacecraft as a function of the time required to accomplish proximity operation maneuvers.
Advisors/Committee Members: David Bradley Spencer, Dissertation Advisor/Co-Advisor, David Bradley Spencer, Committee Chair/Co-Chair, Robert Graham Melton, Committee Member, Sven G Bilen, Committee Member, Joseph Paul Cusumano, Outside Member.
Subjects/Keywords: Astrodynamics; Moon; Mars; Phobos; Proximity Operations; Relative Motion; Semi-Analytical
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APA ·
Chicago ·
MLA ·
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CSE |
Export
to Zotero / EndNote / Reference
Manager
APA (6th Edition):
Conte, D. (2019). Semi-Analytical Solutions for Proximity Operations in the Circular Restricted Three-Body Problem. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/16596dxc5120
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Conte, Davide. “Semi-Analytical Solutions for Proximity Operations in the Circular Restricted Three-Body Problem.” 2019. Thesis, Penn State University. Accessed March 04, 2021.
https://submit-etda.libraries.psu.edu/catalog/16596dxc5120.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Conte, Davide. “Semi-Analytical Solutions for Proximity Operations in the Circular Restricted Three-Body Problem.” 2019. Web. 04 Mar 2021.
Vancouver:
Conte D. Semi-Analytical Solutions for Proximity Operations in the Circular Restricted Three-Body Problem. [Internet] [Thesis]. Penn State University; 2019. [cited 2021 Mar 04].
Available from: https://submit-etda.libraries.psu.edu/catalog/16596dxc5120.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Conte D. Semi-Analytical Solutions for Proximity Operations in the Circular Restricted Three-Body Problem. [Thesis]. Penn State University; 2019. Available from: https://submit-etda.libraries.psu.edu/catalog/16596dxc5120
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Penn State University
28.
Sottile, Brad Joseph.
Evolutionary Computation for Spacecraft Trajectory Optimization.
Degree: 2013, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/19135
► Evolutionary Computation has exploded in use in engineering and the applied sciences. For this thesis, three algorithms – Particle Swarm Optimization (PSO), Bacteria Foraging Optimization…
(more)
▼ Evolutionary Computation has exploded in use in engineering and the applied sciences. For this thesis, three algorithms – Particle Swarm Optimization (PSO), Bacteria Foraging Optimization (BFO) and Covariance Matrix Adaptation Evolution Strategy (CMA-ES) – are compared against each other to solve a classic problem in
astrodynamics, the Hohmann transfer. The role of fixed and varying penalties is explored for each algorithm and compared. Each algorithm was run 1000 times and the performance metrics were compared. PSO using fixed penalties ran with an average central processing unit (CPU) time of 0.138 seconds and yielded a mean error of 1.30% and a median error of 0.48%. Using varying penalties, the algorithm ran with an average CPU time of 0.107 seconds and yielded a mean error of 1.78% and a median error of 0.43%. BFO with fixed penalties had a mean CPU time of 0.655 seconds and yielded a 2.19% mean percent error and 1.91% median percent error. For the varying penalty case, BFO averaged a CPU time of 0.727 seconds, a mean percent error of 0.27% and a median 0.36%. CMA-ES with fixed penalties yielded a mean CPU time of 0.572 seconds, a mean percent error of 0.26% and a median percent error of 0.00%. The varying penalty case for CMA-ES yielded a mean CPU time of 0.582 seconds, a mean percent error of 0.43% and a median percent error of 0.43%. The algorithms all excelled in some areas and had poor performance in others, especially as the penalty case varied. A clear result is that algorithm selection is problem-dependent. Suggestions for future work and applications to other problems are provided.
Advisors/Committee Members: Robert Graham Melton, Thesis Advisor/Co-Advisor.
Subjects/Keywords: Astrodynamics; Evolutionary Algorithms; Evolutionary Computation; Evolutionary Strategies; Optimization; Spacecraft Trajectory
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APA ·
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MLA ·
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CSE |
Export
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APA (6th Edition):
Sottile, B. J. (2013). Evolutionary Computation for Spacecraft Trajectory Optimization. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/19135
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Sottile, Brad Joseph. “Evolutionary Computation for Spacecraft Trajectory Optimization.” 2013. Thesis, Penn State University. Accessed March 04, 2021.
https://submit-etda.libraries.psu.edu/catalog/19135.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Sottile, Brad Joseph. “Evolutionary Computation for Spacecraft Trajectory Optimization.” 2013. Web. 04 Mar 2021.
Vancouver:
Sottile BJ. Evolutionary Computation for Spacecraft Trajectory Optimization. [Internet] [Thesis]. Penn State University; 2013. [cited 2021 Mar 04].
Available from: https://submit-etda.libraries.psu.edu/catalog/19135.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Sottile BJ. Evolutionary Computation for Spacecraft Trajectory Optimization. [Thesis]. Penn State University; 2013. Available from: https://submit-etda.libraries.psu.edu/catalog/19135
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Penn State University
29.
Reiter, Jason A.
Numerical And Analytical Solutions To Rapid Collision Avoidance Maneuvers Constrained By Mission Performance Requirements.
Degree: 2016, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/29469
► Collision avoidance maneuvers to prevent orbital collisions between two catalogued objects are typically planned multiple days in advance. If the warning time is decreased to…
(more)
▼ Collision avoidance maneuvers to prevent orbital collisions between two catalogued objects are typically planned multiple days in advance. If the warning time is decreased to less than half-an-orbit in advance, the problem becomes more complex. Typically, the maneuver (assumed to be impulsive) would be placed at perigee or apogee and oriented in the direction that allows for a fuel-optimal maneuver to be performed well before the predicted collision. Instead, for rapid collision avoidance scenarios, finite burn propagation was applied to determine the thrust duration and direction required to reach a desired minimum collision probability. Determining the thrust time and direction for a wide range of orbits and spacecraft properties results in a semi-analytical solution to the collision avoidance problem anywhere in Low-Earth Orbit. The speed at which this method can be applied makes it valuable when minimal time is available to perform such a maneuver.
For many spacecraft missions, even the slightest change in the orbit of the spacecraft may significantly affect its ability to perform to its required specifications. With the high volume of debris in orbit, debris-creating events could occur with no advanced notice, making rapid collision avoidance scenarios a real possibility. Care must be taken to ensure that any potential collision is avoided while minimizing the effect of the maneuver on the spacecraft’s mission performance. Assuming perfect knowledge of the states of all objects and that the possible collisions occur at high relative velocities, the required thrusting time to achieve a desired collision probability is found. Varying the desired collision probability, the resulting changes in the required thrust duration time (and, thus, fuel use) can be observed, providing options for trading the fuel use and likelihood of a collision. Additionally, both of these variables contribute directly to the ability of the spacecraft to perform to the desired mission specifications. As the collision probability threshold and required burn time increase, the mission performance decreases. The level of robustness necessary in the mission specifications can be used to limit the desired collision probability threshold. This is accomplished by determining the time and fuel required to perform the collision avoidance maneuver to the desired probability level and analyzing the effect of the time spent away from the mission orbit and the quantity of fuel required to perform the maneuver on the mission performance. It was found that, for notification times less than around 20 minutes, it is best to decrease the collision probability as much as the available fuel will allow without regard for the time duration of the maneuver. As the notification time increases past 20 minutes, more emphasis can be placed on the time required to perform the entire maneuver and it was found that simultaneously minimizing the maneuver time and collision probability outweighed the slight extra fuel required for such a maneuver. Such analysis would…
Advisors/Committee Members: David Spencer, Thesis Advisor/Co-Advisor.
Subjects/Keywords: astrodynamics; collision avoidance; debris; trade studies; optimization; trajectory; mission design
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APA ·
Chicago ·
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Export
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Manager
APA (6th Edition):
Reiter, J. A. (2016). Numerical And Analytical Solutions To Rapid Collision Avoidance Maneuvers Constrained By Mission Performance Requirements. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/29469
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Reiter, Jason A. “Numerical And Analytical Solutions To Rapid Collision Avoidance Maneuvers Constrained By Mission Performance Requirements.” 2016. Thesis, Penn State University. Accessed March 04, 2021.
https://submit-etda.libraries.psu.edu/catalog/29469.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Reiter, Jason A. “Numerical And Analytical Solutions To Rapid Collision Avoidance Maneuvers Constrained By Mission Performance Requirements.” 2016. Web. 04 Mar 2021.
Vancouver:
Reiter JA. Numerical And Analytical Solutions To Rapid Collision Avoidance Maneuvers Constrained By Mission Performance Requirements. [Internet] [Thesis]. Penn State University; 2016. [cited 2021 Mar 04].
Available from: https://submit-etda.libraries.psu.edu/catalog/29469.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Reiter JA. Numerical And Analytical Solutions To Rapid Collision Avoidance Maneuvers Constrained By Mission Performance Requirements. [Thesis]. Penn State University; 2016. Available from: https://submit-etda.libraries.psu.edu/catalog/29469
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Penn State University
30.
Kuppa, Koundinya.
Long-term Orbit Propagation Using Symplectic Integration Algorithms.
Degree: 2016, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/29463
► Understanding the evolution of satellite orbits in the long-term is of great importance in astrodynamics. In order to achieve this, accurate propagation of the orbital…
(more)
▼ Understanding the evolution of satellite orbits in the long-term is of great importance in
astrodynamics. In order to achieve this, accurate propagation of the orbital dynamics of the satellite is required. This paper presents implementation and evaluation of a class of numerical integration methods known as symplectic algorithms. This class of algorithms is highly regarded in scientific applications, especially in long-term studies. The objective of this paper is to demonstrate the superior accuracy and efficient speed of several algorithms of this class and obtain long-term state of satellites under the several influencing forces. Within each application, several cases with different values for parameters such as the time step and duration are executed. In addition, long-term orbital evolution of a satellite in various orbital regimes is conducted. The results indicate that the symplectic algorithms are more accurate for orbit propagation at various time increments tested. In addition, the symplectic algorithms are more computationally efficient in all but a few cases.
Advisors/Committee Members: David Bradley Spencer, Thesis Advisor/Co-Advisor.
Subjects/Keywords: astrodynamics; orbital mechanics; symplectic integration; numerical integration; orbit propagation; spaceflight mechanics
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APA ·
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MLA ·
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CSE |
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APA (6th Edition):
Kuppa, K. (2016). Long-term Orbit Propagation Using Symplectic Integration Algorithms. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/29463
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Kuppa, Koundinya. “Long-term Orbit Propagation Using Symplectic Integration Algorithms.” 2016. Thesis, Penn State University. Accessed March 04, 2021.
https://submit-etda.libraries.psu.edu/catalog/29463.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Kuppa, Koundinya. “Long-term Orbit Propagation Using Symplectic Integration Algorithms.” 2016. Web. 04 Mar 2021.
Vancouver:
Kuppa K. Long-term Orbit Propagation Using Symplectic Integration Algorithms. [Internet] [Thesis]. Penn State University; 2016. [cited 2021 Mar 04].
Available from: https://submit-etda.libraries.psu.edu/catalog/29463.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Kuppa K. Long-term Orbit Propagation Using Symplectic Integration Algorithms. [Thesis]. Penn State University; 2016. Available from: https://submit-etda.libraries.psu.edu/catalog/29463
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
◁ [1] [2] [3] [4] [5] ▶
. 
Open Access Theses and Dissertations
