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…
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▼ 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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Iyengar SD. Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model. [Internet] [Masters thesis]. Embry-Riddle Aeronautical University; 2015. [cited 2020 Jan 27].
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
Cal Poly
2.
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Farahmand M. ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK. [Internet] [Masters thesis]. Cal Poly; 2009. [cited 2020 Jan 27].
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
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Sestak TA. The Effect of Surface Materials and Morphology on Wingsuit Aerodynamics. [Internet] [Doctoral dissertation]. Embry-Riddle Aeronautical University; 2017. [cited 2020 Jan 27].
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
4.
Vittaldev, V.
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
Subjects/Keywords: Astrodynamics
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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. “The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.:.” 2010. Masters Thesis, Delft University of Technology. Accessed January 27, 2020.
http://resolver.tudelft.nl/uuid:198390b8-b81f-4a13-9c80-15befd3c0428.
MLA Handbook (7th Edition):
Vittaldev, V. “The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.:.” 2010. Web. 27 Jan 2020.
Vancouver:
Vittaldev V. The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.:. [Internet] [Masters thesis]. Delft University of Technology; 2010. [cited 2020 Jan 27].
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
Purdue University
5.
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 January 27, 2020.
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. 27 Jan 2020.
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 2020 Jan 27].
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
Embry-Riddle Aeronautical University
6.
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…
(more)
▼ 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 January 27, 2020.
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. 27 Jan 2020.
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 2020 Jan 27].
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
Cal Poly
7.
Strange, Michael R.
Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes.
Degree: MS, Aerospace Engineering, 2017, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/1714
;
10.15368/theses.2017.9
► This thesis is directed toward the feasibility of observing satellites on the nano scale and determining an accurate propagated orbit using a Meade LX600-ACF…
(more)
▼ This thesis is directed toward the feasibility of observing satellites on the nano scale and determining an accurate propagated orbit using a Meade LX600-ACF 14” diameter aperture telescope currently located on the California Polytechnic State University campus. The optical telescope is fitted with an f/6.3 focal reducer, SBIG ST-10XME CCD camera and Optec TCF-S Focuser. This instrumentation allowed for a 22’ X 15’ arcminute FOV in order to accurately image passing LEO satellites. Through the use of the Double-r and Gauss Initial Orbit Determination methods as well as Least Squared Differential Correction and Extended Kalman Filter Orbit Determination methods, an accurate predicted orbit can be determined.
These calculated values from observational data of satellites within the Globalstar system are compared against the most updated TLEs for each satellite at the time of observation. The determined differential errors from the well-defined TLEs acquired via online database were used to verify the feasibility of the accuracy which can be obtained from independent observations. Through minimization of error caused from imaging noise, pointing error, and timing error, the main determination of accurate orbital determination lies in the instrumentation mechanical capabilities itself. With the ability to acquire up to 7 individual satellite observations during a single transit, the use of both IOD and OD methods, and the recently acquired Cal Poly telescope with an increased 14” aperture, the feasibility of imaging and orbital determination of nanosatellites is greatly improved.
Advisors/Committee Members: Kira Abercromby.
Subjects/Keywords: observation; satellite; propogation; Astrodynamics
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APA (6th Edition):
Strange, M. R. (2017). Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/1714 ; 10.15368/theses.2017.9
Chicago Manual of Style (16th Edition):
Strange, Michael R. “Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes.” 2017. Masters Thesis, Cal Poly. Accessed January 27, 2020.
https://digitalcommons.calpoly.edu/theses/1714 ; 10.15368/theses.2017.9.
MLA Handbook (7th Edition):
Strange, Michael R. “Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes.” 2017. Web. 27 Jan 2020.
Vancouver:
Strange MR. Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes. [Internet] [Masters thesis]. Cal Poly; 2017. [cited 2020 Jan 27].
Available from: https://digitalcommons.calpoly.edu/theses/1714 ; 10.15368/theses.2017.9.
Council of Science Editors:
Strange MR. Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes. [Masters Thesis]. Cal Poly; 2017. Available from: https://digitalcommons.calpoly.edu/theses/1714 ; 10.15368/theses.2017.9
University of Illinois – Urbana-Champaign
8.
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Ponnapalli K. Optimal variable-specific-impulse trajectories in simple gravitational fields. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2019. [cited 2020 Jan 27].
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
Cal Poly
9.
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Iuliano JR. A Solution to the Circular Restricted N Body Problem in Planetary Systems. [Internet] [Masters thesis]. Cal Poly; 2016. [cited 2020 Jan 27].
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
10.
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Jackson DJ. Formulation of an Optimal Search Strategy for Space Debris At GEO. [Internet] [Masters thesis]. Cal Poly; 2011. [cited 2020 Jan 27].
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
Delft University of Technology
11.
Hess, J.R.
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…
Advisors/Committee Members: Mooij, E., Sudmeijer, K.J..
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. “Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists:.” 2016. Masters Thesis, Delft University of Technology. Accessed January 27, 2020.
http://resolver.tudelft.nl/uuid:dee2d842-ca91-4086-8ef8-771cfee1b0dc.
MLA Handbook (7th Edition):
Hess, J R. “Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists:.” 2016. Web. 27 Jan 2020.
Vancouver:
Hess JR. Aerogravity assists: Hypersonic maneuvering to improve planetary gravity assists:. [Internet] [Masters thesis]. Delft University of Technology; 2016. [cited 2020 Jan 27].
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
University of Colorado
12.
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Geeraert JL. Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications. [Internet] [Doctoral dissertation]. University of Colorado; 2017. [cited 2020 Jan 27].
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
University of New South Wales
13.
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 January 27, 2020.
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. 27 Jan 2020.
Vancouver:
Capon C. Ionospheric Aerodynamics in Low Earth Orbit. [Internet] [Doctoral dissertation]. University of New South Wales; 2017. [cited 2020 Jan 27].
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
Cal Poly
14.
Williams, Adrian Michael.
Thermal Vacuum Chamber Refurbishment and Analysis.
Degree: MS, Aerospace Engineering, 2018, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/1860
;
10.15368/theses.2018.81
► Spacecraft are subject to different environments while on orbit around the Earth and beyond. One of the most critical of these environments that must…
(more)
▼ Spacecraft are
subject to different environments while on orbit around the Earth and beyond. One of the most critical of these environments that must be counteracted is the thermal environment. Each spacecraft has an operating temperature that is specified in the mission requirements. The requirement stems from internal component operating temperatures that are critical to mission success. Prior to placing the spacecraft in orbit, engineers must be sure that the spacecraft will survive or risk losing the mission entirely.
The primary way to mitigate this risk is to use a thermal vacuum chamber (TVAC). The chamber is designed to resemble a space environment by reducing the pressure within the chamber to 1e-6 Torr. The differentiating factor between a vacuum chamber and a thermal vacuum chamber is the ability for the TVAC to complete a process known as thermal cycling using a temperature controller. Thermal cycling begins at a set temperature and increases within the chamber to a designated hot temperature expected to be seen on orbit. After the maximum temperature is reached, it remains there for a specified amount of time in what is called a soak. The controller then reduces the temperature to a specified cold temperature where a second soak takes place. Finally, the temperature is returned to the initial temperature and the process is repeated for a number of cycles until testing is complete.
For the purpose of this thesis, only the initial temperature increase and the first soak are being investigated. The chamber being used to run these experiments was graciously donated by MDA US Systems, however, no additional documentation was provided with the chamber. The Two identical black coated aluminum and brass cylinders have been chosen to be run with three different temperature profiles. The profiles are manually designed in the temperature controller on the chamber and vary by final soak temperature. To supplement the testing, simulations have been created for each test case in order to verify the computer model of the chamber. The simulations utilize AutoCad and Thermal Desktop to provide the results for comparison.
Each of the tests were completed successfully and produced good results that corresponded well to the simulation. The largest difference between the simulation cylinder temperature and the experimental cylinder temperature was 1.9
∘C. The effectiveness and efficiency of the blue chamber was compared to the other thermal vacuum chamber in the Space Environments lab. Overall, the Blue Chamber proved to be more robust and much easier to operate than the HVEC thermal vacuum chamber.
Advisors/Committee Members: Kira Abercromby.
Subjects/Keywords: thermal; vacuum; chamber; space; environments; pressure; Astrodynamics
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APA (6th Edition):
Williams, A. M. (2018). Thermal Vacuum Chamber Refurbishment and Analysis. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/1860 ; 10.15368/theses.2018.81
Chicago Manual of Style (16th Edition):
Williams, Adrian Michael. “Thermal Vacuum Chamber Refurbishment and Analysis.” 2018. Masters Thesis, Cal Poly. Accessed January 27, 2020.
https://digitalcommons.calpoly.edu/theses/1860 ; 10.15368/theses.2018.81.
MLA Handbook (7th Edition):
Williams, Adrian Michael. “Thermal Vacuum Chamber Refurbishment and Analysis.” 2018. Web. 27 Jan 2020.
Vancouver:
Williams AM. Thermal Vacuum Chamber Refurbishment and Analysis. [Internet] [Masters thesis]. Cal Poly; 2018. [cited 2020 Jan 27].
Available from: https://digitalcommons.calpoly.edu/theses/1860 ; 10.15368/theses.2018.81.
Council of Science Editors:
Williams AM. Thermal Vacuum Chamber Refurbishment and Analysis. [Masters Thesis]. Cal Poly; 2018. Available from: https://digitalcommons.calpoly.edu/theses/1860 ; 10.15368/theses.2018.81
University of Colorado
15.
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 January 27, 2020.
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. 27 Jan 2020.
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 2020 Jan 27].
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
16.
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 January 27, 2020.
https://scholar.colorado.edu/asen_gradetds/219.
MLA Handbook (7th Edition):
Baresi, Nicola. “Spacecraft Formation Flight on Quasi-Periodic Invariant Tori.” 2017. Web. 27 Jan 2020.
Vancouver:
Baresi N. Spacecraft Formation Flight on Quasi-Periodic Invariant Tori. [Internet] [Doctoral dissertation]. University of Colorado; 2017. [cited 2020 Jan 27].
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
17.
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 January 27, 2020.
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. 27 Jan 2020.
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 2020 Jan 27].
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
The Ohio State University
18.
Poli, C.
A study of the effect of man's motion on the attitude and
orbital motion of a satellite.
Degree: PhD, Graduate School, 1965, The Ohio State University
URL: http://rave.ohiolink.edu/etdc/view?acc_num=osu1486567958464808
Subjects/Keywords: Education; Artificial satellites; Astrodynamics
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APA (6th Edition):
Poli, C. (1965). A study of the effect of man's motion on the attitude and
orbital motion of a satellite. (Doctoral Dissertation). The Ohio State University. Retrieved from http://rave.ohiolink.edu/etdc/view?acc_num=osu1486567958464808
Chicago Manual of Style (16th Edition):
Poli, C. “A study of the effect of man's motion on the attitude and
orbital motion of a satellite.” 1965. Doctoral Dissertation, The Ohio State University. Accessed January 27, 2020.
http://rave.ohiolink.edu/etdc/view?acc_num=osu1486567958464808.
MLA Handbook (7th Edition):
Poli, C. “A study of the effect of man's motion on the attitude and
orbital motion of a satellite.” 1965. Web. 27 Jan 2020.
Vancouver:
Poli C. A study of the effect of man's motion on the attitude and
orbital motion of a satellite. [Internet] [Doctoral dissertation]. The Ohio State University; 1965. [cited 2020 Jan 27].
Available from: http://rave.ohiolink.edu/etdc/view?acc_num=osu1486567958464808.
Council of Science Editors:
Poli C. A study of the effect of man's motion on the attitude and
orbital motion of a satellite. [Doctoral Dissertation]. The Ohio State University; 1965. Available from: http://rave.ohiolink.edu/etdc/view?acc_num=osu1486567958464808
Texas A&M University
19.
Johnson, Kirk Wayne.
Approaches for Modeling Satellite Relative Motion.
Degree: 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 ?. 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. (Thesis). Texas A&M University. Retrieved from http://hdl.handle.net/1969.1/158945
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):
Johnson, Kirk Wayne. “Approaches for Modeling Satellite Relative Motion.” 2016. Thesis, Texas A&M University. Accessed January 27, 2020.
http://hdl.handle.net/1969.1/158945.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Johnson, Kirk Wayne. “Approaches for Modeling Satellite Relative Motion.” 2016. Web. 27 Jan 2020.
Vancouver:
Johnson KW. Approaches for Modeling Satellite Relative Motion. [Internet] [Thesis]. Texas A&M University; 2016. [cited 2020 Jan 27].
Available from: http://hdl.handle.net/1969.1/158945.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Johnson KW. Approaches for Modeling Satellite Relative Motion. [Thesis]. Texas A&M University; 2016. Available from: http://hdl.handle.net/1969.1/158945
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Illinois – Urbana-Champaign
20.
De La Mata, Jorge.
Optimal strategies for selecting from multiple asteroid interception targets.
Degree: MS, 4048, 2014, University of Illinois – Urbana-Champaign
URL: http://hdl.handle.net/2142/50647
► The goal of this thesis is to demonstrate the applicability of the discrimination time optimization problem to space systems. The concept is applied to an…
(more)
▼ The goal of this thesis is to demonstrate the applicability of the discrimination time optimization problem to space systems. The concept is applied to an asteroid interception problem using low-thrust propulsion and solved through a combination of particle swarm optimization (PSO) and a commercial non-linear programming (NLP) solver SNOPT (Sparse Nonlinear OPTimizer).
The problem considers an intercepting vehicle and three bodies of interest, only one of which can be intercepted. The resulting optimal trajectory must maximize the amount of time the vehicle can travel before committing to intercepting one of the three bodies of interest, assuring that all three bodies are available for intercept within a maximum flight time. This time before commitment is called the discrimination time. Practically, an increase in discrimination time provides the decision makers with an opportunity for further deliberation and organization of mission priorities. This can represent additional time to determine which body holds the most scientific value, or which body presents the largest threat to human life or financial assets.
The solution method is a two-step process. An approximate solution is found using a metaheursitic method, PSO. The control is modeled as a combination of polynomials to give the best opportunity for convergence to a feasible solution. The solution found by the PSO algorithm is then used to initialize a much more accurate solution method using NLP.
The two optimization techniques are applied in series because of the problem complexity. The PSO algorithm is a stochastic process and can therefore not guarantee optimality and struggles to perform with a large set of optimal parameters. NLP solvers can guarantee first or second order optimality, but require initialization and convergence is contingent on the quality of the initial guess. Creating a four phase trajectory solution ad hoc that could reliably serve as an initial guess for the NLP solver is impractical, and allowing the PSO algorithm to solve a fully discrete control time history problem would be impractical. Using the two together takes advantage of the simplicity and speed provided by the PSO algorithm and the ability of the NLP solver to yield high fidelity, optimal solutions.
Advisors/Committee Members: Conway, Bruce A. (advisor).
Subjects/Keywords: Astrodynamics; Orbital Mechanics; Optimization; Nonlinear Programming; Space; Aerospace; Optimal trajectories; Controls
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Export
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APA (6th Edition):
De La Mata, J. (2014). Optimal strategies for selecting from multiple asteroid interception targets. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/50647
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):
De La Mata, Jorge. “Optimal strategies for selecting from multiple asteroid interception targets.” 2014. Thesis, University of Illinois – Urbana-Champaign. Accessed January 27, 2020.
http://hdl.handle.net/2142/50647.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
De La Mata, Jorge. “Optimal strategies for selecting from multiple asteroid interception targets.” 2014. Web. 27 Jan 2020.
Vancouver:
De La Mata J. Optimal strategies for selecting from multiple asteroid interception targets. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2014. [cited 2020 Jan 27].
Available from: http://hdl.handle.net/2142/50647.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
De La Mata J. Optimal strategies for selecting from multiple asteroid interception targets. [Thesis]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/50647
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Penn State University
21.
Reiter, Jason A.
Numerical And Analytical Solutions To Rapid Collision
Avoidance Maneuvers Constrained By Mission Performance
Requirements.
Degree: MS, Aerospace Engineering, 2016, Penn State University
URL: https://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…
Subjects/Keywords: astrodynamics; collision avoidance; debris; trade studies;
optimization; trajectory; mission design
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APA ·
Chicago ·
MLA ·
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CSE |
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. (Masters Thesis). Penn State University. Retrieved from https://etda.libraries.psu.edu/catalog/29469
Chicago Manual of Style (16th Edition):
Reiter, Jason A. “Numerical And Analytical Solutions To Rapid Collision
Avoidance Maneuvers Constrained By Mission Performance
Requirements.” 2016. Masters Thesis, Penn State University. Accessed January 27, 2020.
https://etda.libraries.psu.edu/catalog/29469.
MLA Handbook (7th Edition):
Reiter, Jason A. “Numerical And Analytical Solutions To Rapid Collision
Avoidance Maneuvers Constrained By Mission Performance
Requirements.” 2016. Web. 27 Jan 2020.
Vancouver:
Reiter JA. Numerical And Analytical Solutions To Rapid Collision
Avoidance Maneuvers Constrained By Mission Performance
Requirements. [Internet] [Masters thesis]. Penn State University; 2016. [cited 2020 Jan 27].
Available from: https://etda.libraries.psu.edu/catalog/29469.
Council of Science Editors:
Reiter JA. Numerical And Analytical Solutions To Rapid Collision
Avoidance Maneuvers Constrained By Mission Performance
Requirements. [Masters Thesis]. Penn State University; 2016. Available from: https://etda.libraries.psu.edu/catalog/29469
Penn State University
22.
Scott, Christopher James.
Transfer and Capture into Distant Retrograde Orbits.
Degree: PhD, Aerospace Engineering, 2010, Penn State University
URL: https://etda.libraries.psu.edu/catalog/10471
► This dissertation utilizes theory and techniques derived from the fields of dynamical systems theory, astrodyanamics, celestial mechanics, and fluid mechanics to analyze the phenomenon of…
(more)
▼ This dissertation utilizes theory and techniques
derived from the fields of dynamical systems theory,
astrodyanamics, celestial mechanics, and fluid mechanics to analyze
the phenomenon of satellite capture and interrelated spacecraft
transfers in restricted three-body systems. The results extend
current knowledge and understanding of capture dynamics in the
context of astrodynamics and celestial mechanics. Manifold theory,
fast Lyapunov indicator maps, and the classification of space
structure facilitate an analysis of the transport of objects from
the chaotic reaches of the solar system to the distant retrograde
region in the sun-Jupiter system. Apart from past studies this
dissertation considers the role of the complex lobe structure
encompassing stable regions in the circular restricted three-body
problem. These structures are shown to be responsible for the
phenomenon of sticky orbits and the transport of objects among
stable regions. Since permanent capture can only be achieved
through a change in energy, fast Lyapunov indicator maps and other
methods which reveal the structure of the conservative system are
used to discern capture regions and identify the underpinnings of
the dynamics. Fast Lyapunov indicator maps provide an accurate
classification of orbits of permanent capture and escape, yet
monopolize computational resources. In anticipation of a fully
three-dimensional analysis in the dissipative system a new mapping
parameter is introduced based on energy degradation and averaged
velocity. Although the study specifically addresses the sun-Jupiter
system, the qualitative results and devised techniques can be
applied throughout the solar system and to capture about extrasolar
planets. Extending the analysis beyond the exterior of the stable
distant retrograde region fosters the construction of transfer
orbits from low-Earth orbit to a stable periodic orbit at the
center of the stable distant retrograde region. Key to this
analysis is the predictability of collision orbits within the
highly chaotic region commonly recognized as a saddle point on the
energy manifold. The pragmatic techniques derived from this
analysis solve a number of complications apparent in the
literature. Notably a reliable methodology for the construction of
an arbitrary number of transfer orbits circumvents the requirement
of computing specialized periodic orbits or extensive numerical
sampling of the phase space. The procedure provides a complete
description of the design space accessing a wide range of distant
retrograde orbits sizes, insertion points, and parking orbit
altitudes in an automated manner. The transfers are studied in a
similar fashion to periodic orbits unveiling the intimate
relationship among design parameters and phase space structure. An
arbitrary number of Earth return periodic orbits can be generated
as a by-product. These orbits may be useful for spacecraft that
must make a number of passes near the second primary without a
reduction in energy. Further analysis of the lobe dynamics and a
modification of the…
Subjects/Keywords: circular restricted three body problem; Astrodynamics;
CR3BP; satellite capture
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APA ·
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MLA ·
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CSE |
Export
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APA (6th Edition):
Scott, C. J. (2010). Transfer and Capture into Distant Retrograde Orbits. (Doctoral Dissertation). Penn State University. Retrieved from https://etda.libraries.psu.edu/catalog/10471
Chicago Manual of Style (16th Edition):
Scott, Christopher James. “Transfer and Capture into Distant Retrograde Orbits.” 2010. Doctoral Dissertation, Penn State University. Accessed January 27, 2020.
https://etda.libraries.psu.edu/catalog/10471.
MLA Handbook (7th Edition):
Scott, Christopher James. “Transfer and Capture into Distant Retrograde Orbits.” 2010. Web. 27 Jan 2020.
Vancouver:
Scott CJ. Transfer and Capture into Distant Retrograde Orbits. [Internet] [Doctoral dissertation]. Penn State University; 2010. [cited 2020 Jan 27].
Available from: https://etda.libraries.psu.edu/catalog/10471.
Council of Science Editors:
Scott CJ. Transfer and Capture into Distant Retrograde Orbits. [Doctoral Dissertation]. Penn State University; 2010. Available from: https://etda.libraries.psu.edu/catalog/10471
Penn State University
23.
Kuppa, Koundinya.
Long-term Orbit Propagation Using Symplectic Integration
Algorithms.
Degree: MS, Aerospace Engineering, 2016, Penn State University
URL: https://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.
Subjects/Keywords: astrodynamics; orbital mechanics; symplectic integration;
numerical integration; orbit propagation; spaceflight
mechanics
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APA ·
Chicago ·
MLA ·
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CSE |
Export
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Manager
APA (6th Edition):
Kuppa, K. (2016). Long-term Orbit Propagation Using Symplectic Integration
Algorithms. (Masters Thesis). Penn State University. Retrieved from https://etda.libraries.psu.edu/catalog/29463
Chicago Manual of Style (16th Edition):
Kuppa, Koundinya. “Long-term Orbit Propagation Using Symplectic Integration
Algorithms.” 2016. Masters Thesis, Penn State University. Accessed January 27, 2020.
https://etda.libraries.psu.edu/catalog/29463.
MLA Handbook (7th Edition):
Kuppa, Koundinya. “Long-term Orbit Propagation Using Symplectic Integration
Algorithms.” 2016. Web. 27 Jan 2020.
Vancouver:
Kuppa K. Long-term Orbit Propagation Using Symplectic Integration
Algorithms. [Internet] [Masters thesis]. Penn State University; 2016. [cited 2020 Jan 27].
Available from: https://etda.libraries.psu.edu/catalog/29463.
Council of Science Editors:
Kuppa K. Long-term Orbit Propagation Using Symplectic Integration
Algorithms. [Masters Thesis]. Penn State University; 2016. Available from: https://etda.libraries.psu.edu/catalog/29463
Penn State University
24.
Sottile, Bradley Joseph.
Evolutionary Computation for Spacecraft Trajectory
Optimization.
Degree: MS, Aerospace Engineering, 2013, Penn State University
URL: https://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.
Subjects/Keywords: Astrodynamics; Evolutionary Algorithms; Evolutionary
Computation; Evolutionary Strategies; Optimization; Spacecraft
Trajectory
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APA ·
Chicago ·
MLA ·
Vancouver ·
CSE |
Export
to Zotero / EndNote / Reference
Manager
APA (6th Edition):
Sottile, B. J. (2013). Evolutionary Computation for Spacecraft Trajectory
Optimization. (Masters Thesis). Penn State University. Retrieved from https://etda.libraries.psu.edu/catalog/19135
Chicago Manual of Style (16th Edition):
Sottile, Bradley Joseph. “Evolutionary Computation for Spacecraft Trajectory
Optimization.” 2013. Masters Thesis, Penn State University. Accessed January 27, 2020.
https://etda.libraries.psu.edu/catalog/19135.
MLA Handbook (7th Edition):
Sottile, Bradley Joseph. “Evolutionary Computation for Spacecraft Trajectory
Optimization.” 2013. Web. 27 Jan 2020.
Vancouver:
Sottile BJ. Evolutionary Computation for Spacecraft Trajectory
Optimization. [Internet] [Masters thesis]. Penn State University; 2013. [cited 2020 Jan 27].
Available from: https://etda.libraries.psu.edu/catalog/19135.
Council of Science Editors:
Sottile BJ. Evolutionary Computation for Spacecraft Trajectory
Optimization. [Masters Thesis]. Penn State University; 2013. Available from: https://etda.libraries.psu.edu/catalog/19135
University of Southern California
25.
Sarzi Amade', Nicola.
Mars rapid round trip mission design.
Degree: PhD, Astronautical Engineering, 2010, University of Southern California
URL: http://digitallibrary.usc.edu/cdm/compoundobject/collection/p15799coll127/id/361721/rec/3962
► The present research is divided in two parts. The first part is a well defined mathematical problem, with exact rules and results, in which the…
(more)
▼ The present research is divided in two parts. The
first part is a well defined mathematical problem, with exact rules
and results, in which the basic constraints for interplanetary
round trip travels are used to calculate an interplanetary train
schedule (ITS) of missions to Mars, in the general case of orbits
with non-zero eccentricity and non-zero inclination. Several
possible options for round trip travels to Mars are considered. In
particular, options at high energy, which allow rapid round trip
missions, are discussed. These options have important applications
for human travels to Mars.; The second part of the research is
about systems engineering aspects, which are intrinsically less
exact, since they can change with time due, for example, to
technology development or economic and political factors. For the
case of a selected human rapid round trip mission to Mars, the
development of a mission architecture, an assessment of the masses
involved in the mission (such as the initial masses required in
LEO), an estimate of the necessary number of launches, and a
preliminary analysis of the radiation protection requirements, are
performed.; The main problem that justifies the existence of basic
constraints for round trip missions is that by increasing the ΔV of
a mission, in general the total round trip time does not vary much,
because a higher ΔV can only reduce the transfer time and it simply
increases the stay-time on the target planet. However, if the ΔV is
increased beyond a well-defined level, the total round trip time
has a sudden drop in duration that makes fast round trips possible.
This is due to the fact that the traveler can go back before the
home planet makes one extra revolution around the Sun. For a
sufficiently high ΔV, a round trip to Mars can change in duration
from 2.7 years to about 5 months.; For Mars missions, the round
trip times are calculated for different ΔV’s and for different
transfer trajectories (T1, T2, etc.). An interplanetary train
schedule is created. The behavior of the constraints near transfer
time limits, and the impact of real vs. ideal orbital elements of
the Earth and Mars (where “ideal” indicates circular and coplanar
orbits) are analyzed.; It was found that for the profiles examined
(T1/T1, T1/T4, and NT1/NT4) and for the opposition of July 27th,
2018, the transfer times can be shorter by up to 4 weeks at
outbound and by up to 6 weeks at return, compared to circular
orbits. Instead, for the opposition of February 19th, 2027, the
transfer times can be longer by up to 6 weeks both at outbound and
return, compared to circular orbits. Also, for the opposition of
February 19th, 2027, rapid T1/T1 round trip missions are not
possible given the farther distance between the Earth and Mars. In
that case, only long-stay missions are possible.; For the systems
engineering part of the research, a comparison between the
requirements of a representative rapid round trip mission and those
of three NASA Mars reference missions is made, in order to
understand the relative advantages and…
Advisors/Committee Members: Erwin, Daniel A. (Committee Chair), Wertz, James R. (Committee Member), Gruntman, Mike (Committee Member), Udwadia, Firdaus E. (Committee Member).
Subjects/Keywords: Mars; astrodynamics; roundtrip; rapid roundtrip; human spaceflight; systems engineering; mission design
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APA (6th Edition):
Sarzi Amade', N. (2010). Mars rapid round trip mission design. (Doctoral Dissertation). University of Southern California. Retrieved from http://digitallibrary.usc.edu/cdm/compoundobject/collection/p15799coll127/id/361721/rec/3962
Chicago Manual of Style (16th Edition):
Sarzi Amade', Nicola. “Mars rapid round trip mission design.” 2010. Doctoral Dissertation, University of Southern California. Accessed January 27, 2020.
http://digitallibrary.usc.edu/cdm/compoundobject/collection/p15799coll127/id/361721/rec/3962.
MLA Handbook (7th Edition):
Sarzi Amade', Nicola. “Mars rapid round trip mission design.” 2010. Web. 27 Jan 2020.
Vancouver:
Sarzi Amade' N. Mars rapid round trip mission design. [Internet] [Doctoral dissertation]. University of Southern California; 2010. [cited 2020 Jan 27].
Available from: http://digitallibrary.usc.edu/cdm/compoundobject/collection/p15799coll127/id/361721/rec/3962.
Council of Science Editors:
Sarzi Amade' N. Mars rapid round trip mission design. [Doctoral Dissertation]. University of Southern California; 2010. Available from: http://digitallibrary.usc.edu/cdm/compoundobject/collection/p15799coll127/id/361721/rec/3962
26.
Conte, Davide.
Determination of Optimal Earth-mars Trajectories to Target
the Moons of Mars.
Degree: MS, Aerospace Engineering, 2014, Penn State University
URL: https://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.…
(more)
▼ 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.
Subjects/Keywords: Mars; Phobos; Deimos; astrodynamics
…undergraduate and graduate theses, and especially for teaching Interplanetary Astrodynamics, one
of… …in the field of astrodynamics that the Swiss mathematician Johann Heinrich Lambert… …astrodynamics, Lambert’s
problem is a two-point boundary value problem (see Figure 8) for… …eccentricity. Such derivation can be found in most astrodynamics
textbooks such as Prussing and…
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Manager
APA (6th Edition):
Conte, D. (2014). Determination of Optimal Earth-mars Trajectories to Target
the Moons of Mars. (Masters Thesis). Penn State University. Retrieved from https://etda.libraries.psu.edu/catalog/21382
Chicago Manual of Style (16th Edition):
Conte, Davide. “Determination of Optimal Earth-mars Trajectories to Target
the Moons of Mars.” 2014. Masters Thesis, Penn State University. Accessed January 27, 2020.
https://etda.libraries.psu.edu/catalog/21382.
MLA Handbook (7th Edition):
Conte, Davide. “Determination of Optimal Earth-mars Trajectories to Target
the Moons of Mars.” 2014. Web. 27 Jan 2020.
Vancouver:
Conte D. Determination of Optimal Earth-mars Trajectories to Target
the Moons of Mars. [Internet] [Masters thesis]. Penn State University; 2014. [cited 2020 Jan 27].
Available from: https://etda.libraries.psu.edu/catalog/21382.
Council of Science Editors:
Conte D. Determination of Optimal Earth-mars Trajectories to Target
the Moons of Mars. [Masters Thesis]. Penn State University; 2014. Available from: https://etda.libraries.psu.edu/catalog/21382
Cal Poly
27.
Smiroldo, Jordan.
Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update.
Degree: MS, Aerospace Engineering, 2013, Cal Poly
URL: https://digitalcommons.calpoly.edu/theses/1135
;
10.15368/theses.2013.215
► The renewed academic interest in using solar sails as a source of spacecraft propulsion has been accompanied by a recent fervor of investigations into…
(more)
▼ The renewed academic interest in using solar sails as a source of spacecraft propulsion has been accompanied by a recent fervor of investigations into non-ideal and off-nominal sail performance considerations. One of the most influential considerations, uncertain optical degradation, has been shown to present significant trajectory design difficulties. This paper investigates the potential of using a mid-course degradation model update to mitigate the risk of missing the target destination in a sample 300 day Earth-Venus trajectory. Using a range of potential degradation profiles, it is shown that correcting in the first half of the mission is highly likely to result in a trajectory that arrives sufficiently close to Venus at the end of the mission timeframe. Depending on the exact extent of the uncertainty, the data suggests that the latest a correction should take place ranges from 150 to 240 days into the mission. The influence of two different parameters, the extent and rate of degradation, are compared to show that the former of the two is more impactful on correcting timing than the latter.
Advisors/Committee Members: Dr. Kira Abercromby.
Subjects/Keywords: Solar; Sail; Trajectory; Collocation; Non-ideal; Performance; Astrodynamics
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APA (6th Edition):
Smiroldo, J. (2013). Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update. (Masters Thesis). Cal Poly. Retrieved from https://digitalcommons.calpoly.edu/theses/1135 ; 10.15368/theses.2013.215
Chicago Manual of Style (16th Edition):
Smiroldo, Jordan. “Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update.” 2013. Masters Thesis, Cal Poly. Accessed January 27, 2020.
https://digitalcommons.calpoly.edu/theses/1135 ; 10.15368/theses.2013.215.
MLA Handbook (7th Edition):
Smiroldo, Jordan. “Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update.” 2013. Web. 27 Jan 2020.
Vancouver:
Smiroldo J. Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update. [Internet] [Masters thesis]. Cal Poly; 2013. [cited 2020 Jan 27].
Available from: https://digitalcommons.calpoly.edu/theses/1135 ; 10.15368/theses.2013.215.
Council of Science Editors:
Smiroldo J. Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update. [Masters Thesis]. Cal Poly; 2013. Available from: https://digitalcommons.calpoly.edu/theses/1135 ; 10.15368/theses.2013.215
Virginia Tech
28.
Rogers, Andrew Charles.
Optimization-Based Guidance for Satellite Relative Motion.
Degree: PhD, Aerospace and Ocean Engineering, 2016, Virginia Tech
URL: http://hdl.handle.net/10919/79455
► Spacecraft relative motion modeling and control promises to enable or augment a wide range of missions for scientific research, military applications, and space situational awareness.…
(more)
▼ Spacecraft relative motion modeling and control promises to enable or augment a wide range of missions for scientific research, military applications, and space situational awareness. This dissertation focuses on the development of novel, optimization-based, control design for some representative relative-motion-enabled missions. Spacecraft relative motion refers to two (or more) satellites in nearly identical orbits. We examine control design for relative configurations on the scale of meters (for the purposes of proximity operations) as well as on the scale of tens of kilometers (representative of science gathering missions). Realistic control design for satellites is limited by accurate modeling of the relative orbital perturbations as well as the highly constrained nature of most space systems. We present solutions to several types of optimal orbital maneuvers using a variety of different, realistic assumptions based on the maneuver objectives.
Initially, we assume a perfectly circular orbit with a perfectly spherical Earth and analytically solve the under-actuated, minimum-energy, optimal transfer using techniques from optimal control and linear operator theory. The resulting open-loop control law is guaranteed to be a global optimum. Then, recognizing that very few, if any, orbits are truly circular, the optimal transfer problem is generalized to the elliptical linear and nonlinear systems which describe the relative motion. Solution of the minimum energy transfer for both the linear and nonlinear systems reveals that the resulting trajectories are nearly identical, implying that the nonlinearity has little effect on the relative motion. A continuous-time, nonlinear, sliding mode controller which tracks the linear trajectory in the presence of a higher fidelity orbit model shows that the closed-loop system is both asymptotically stable and robust to disturbances and un-modeled dynamics.
Next, a novel method of computing discrete-time, multi-revolution, finite-thrust, fuel-optimal, relative orbit transfers near an elliptical, perturbed orbit is presented. The optimal control problem is based on the classical, continuous-time, fuel-optimization problem from calculus of variations, and we present the discrete-time analogue of this problem using a transcription-based method. The resulting linear program guarantees a global optimum in terms of fuel consumption, and we validate the results using classical impulsive orbit transfer theory. The new method is shown to converge to classical impulsive orbit transfer theory in the limit that the duration of the zero-order hold discretization approaches zero and the time horizon extends to infinity. Then the fuel/time optimal control problem is solved using a hybrid approach which uses a linear program to solve the fuel optimization, and a genetic algorithm to find the minimizing time-of-flight. The method developed in this work allows mission planners to determine the feasibility for realistic spacecraft and motion models.
Proximity operations for robotic…
Advisors/Committee Members: Woolsey, Craig A. (committeechair), McGwier, Robert W. (committeechair), Ross, Shane David (committee member), Earle, Gregory D. (committee member), Black, Jonathan T. (committee member).
Subjects/Keywords: Model predictive control; astrodynamics; optimal control; optimization; proximity operations
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APA (6th Edition):
Rogers, A. C. (2016). Optimization-Based Guidance for Satellite Relative Motion. (Doctoral Dissertation). Virginia Tech. Retrieved from http://hdl.handle.net/10919/79455
Chicago Manual of Style (16th Edition):
Rogers, Andrew Charles. “Optimization-Based Guidance for Satellite Relative Motion.” 2016. Doctoral Dissertation, Virginia Tech. Accessed January 27, 2020.
http://hdl.handle.net/10919/79455.
MLA Handbook (7th Edition):
Rogers, Andrew Charles. “Optimization-Based Guidance for Satellite Relative Motion.” 2016. Web. 27 Jan 2020.
Vancouver:
Rogers AC. Optimization-Based Guidance for Satellite Relative Motion. [Internet] [Doctoral dissertation]. Virginia Tech; 2016. [cited 2020 Jan 27].
Available from: http://hdl.handle.net/10919/79455.
Council of Science Editors:
Rogers AC. Optimization-Based Guidance for Satellite Relative Motion. [Doctoral Dissertation]. Virginia Tech; 2016. Available from: http://hdl.handle.net/10919/79455
Delft University of Technology
29.
Hoogland, J.
Earth frozen orbits: design, injection and stability:.
Degree: 2016, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:fc0a57ce-33df-4cd2-a61b-66e702cc9cf8
► A frozen orbit is an orbit chosen such that the effect of perturbations on (a combination of) the mean orbital elements is minimized. The concept…
(more)
▼ A frozen orbit is an orbit chosen such that the effect of perturbations on (a combination of) the mean orbital elements is minimized. The concept first appeared in literature in 1978, and was applied that same year to the Seasat mission. This altimetry mission featured strict requirements on the accuracy of the altitude of the satellite above the sea surface. By designing an orbit for which the mean eccentricity and mean argument of periapsis remain static, the satellite’s altitude will theoretically be constant, depending only on location of the sub-satellite point.
Classically, the theory behind frozen orbits is only based on the J2- and J3-term of the spherical har- monics gravity field model and clever manipulation of the Lagrange planetary equations. Through considerable analytical effort, it is possible to include all other zonal gravity field terms into the equation, but this approach is limited to perturbations that can be cast into the form of a disturbing potential. The aim of this thesis is to find a numerical method that overcomes this limit and to use that method to investigate the effects of including third-body gravity, atmospheric drag and solar radiation pressure on the mean orbital elements.
To do this, the frozen orbit problem is formulated as an optimization problem. Use is made of Differential Evolution (DE) and grid searching to simulate many trajectories and to find a set of injection parameters that results in a minimal variation in the mean eccentricity and mean argument of periapsis. The mean elements are reconstructed from the osculating elements by making use of the Eckstein-Ustinov theory and subsequent numerical averaging. In combination with Precise Orbit Determination (POD) data, this reconstruction is used to investigate the variations in the mean orbital elements of ERS-2 and TOPEX/Poseidon.
Subsequently, the numerical method is applied to various orbital dynamics models. When applied to zonal gravity fields, the new method is found to be in good agreement with analytical solutions. The influence of other perturbations on solutions found in zonal models is examined, and it is found that taking these perturbations into account during the optimization process does not lead to significant improvements with respect to the simple zonal case, nor does it lead to significant changes in the found injection conditions. For the assumed satellite characteristics, radiation pressure is found to be the most influential perturbation, causing fluctuations in the mean eccentricity of ±3%.
Advisors/Committee Members: Noomen, R..
Subjects/Keywords: astrodynamics; frozen orbit; orbital perturbations; mission design; optimization; orbit injection
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APA ·
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MLA ·
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Export
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APA (6th Edition):
Hoogland, J. (2016). Earth frozen orbits: design, injection and stability:. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:fc0a57ce-33df-4cd2-a61b-66e702cc9cf8
Chicago Manual of Style (16th Edition):
Hoogland, J. “Earth frozen orbits: design, injection and stability:.” 2016. Masters Thesis, Delft University of Technology. Accessed January 27, 2020.
http://resolver.tudelft.nl/uuid:fc0a57ce-33df-4cd2-a61b-66e702cc9cf8.
MLA Handbook (7th Edition):
Hoogland, J. “Earth frozen orbits: design, injection and stability:.” 2016. Web. 27 Jan 2020.
Vancouver:
Hoogland J. Earth frozen orbits: design, injection and stability:. [Internet] [Masters thesis]. Delft University of Technology; 2016. [cited 2020 Jan 27].
Available from: http://resolver.tudelft.nl/uuid:fc0a57ce-33df-4cd2-a61b-66e702cc9cf8.
Council of Science Editors:
Hoogland J. Earth frozen orbits: design, injection and stability:. [Masters Thesis]. Delft University of Technology; 2016. Available from: http://resolver.tudelft.nl/uuid:fc0a57ce-33df-4cd2-a61b-66e702cc9cf8
Delft University of Technology
30.
Verzijl, C.J.O.
On the determination of approximations of first integrals for few-body gravitational problems: with applications to capture trajectories:.
Degree: Electrical Engineering, Mathematics and Computer Science, Mathematical Physics, 2007, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:dacac69b-b5f8-4216-a143-825a0e3bc2ea
► On the determination of approximations of first integrals using the method of integrating vectors for ODE systems, as applied to few-body gravitational problems. Considers the…
(more)
▼ On the determination of approximations of first integrals using the method of integrating vectors for ODE systems, as applied to few-body gravitational problems. Considers the Jacobi 3-body problem, the circular restricted 3-body problem and a 4-body model for ballistic lunar capture. Also discusses the application of these techniques to numerical solutions of the ODE systems using methods designed to preserve exact and approximate first integrals, such as those developed using the method of integrating vectors.
Advisors/Committee Members: Van Horssen, W.T., Noomen, R..
Subjects/Keywords: integral; conservative; approximation; astrodynamics; gravitational
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APA ·
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APA (6th Edition):
Verzijl, C. J. O. (2007). On the determination of approximations of first integrals for few-body gravitational problems: with applications to capture trajectories:. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:dacac69b-b5f8-4216-a143-825a0e3bc2ea
Chicago Manual of Style (16th Edition):
Verzijl, C J O. “On the determination of approximations of first integrals for few-body gravitational problems: with applications to capture trajectories:.” 2007. Masters Thesis, Delft University of Technology. Accessed January 27, 2020.
http://resolver.tudelft.nl/uuid:dacac69b-b5f8-4216-a143-825a0e3bc2ea.
MLA Handbook (7th Edition):
Verzijl, C J O. “On the determination of approximations of first integrals for few-body gravitational problems: with applications to capture trajectories:.” 2007. Web. 27 Jan 2020.
Vancouver:
Verzijl CJO. On the determination of approximations of first integrals for few-body gravitational problems: with applications to capture trajectories:. [Internet] [Masters thesis]. Delft University of Technology; 2007. [cited 2020 Jan 27].
Available from: http://resolver.tudelft.nl/uuid:dacac69b-b5f8-4216-a143-825a0e3bc2ea.
Council of Science Editors:
Verzijl CJO. On the determination of approximations of first integrals for few-body gravitational problems: with applications to capture trajectories:. [Masters Thesis]. Delft University of Technology; 2007. Available from: http://resolver.tudelft.nl/uuid:dacac69b-b5f8-4216-a143-825a0e3bc2ea
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Open Access Theses and Dissertations
