Maintaining Space Domain Awareness (SDA) of satellites in low Earth orbit (LEO) requires effective methods of tracking and characterization. Optical measurements of these objects are generally sparse due to limited access intervals and high angular rates. Light pollution and geographic obstructions may also preclude consistent observations. However, a mobile small aperture telescope grants the ability to minimize such environmental effects, thereby increasing capture likelihoods for objects within this regime. By enhancing LEO satellite visibility in this way, extensive orbital and visual data are obtainable. An 8-inch Meade LX200GPS telescope equipped with a Lumenera SKYnyx2-0M CCD camera comprises the system that conducted observations of LEO. From 22 sessions spanning four months, 76 objects were imaged to provide a data set of 313 streak frames for initial orbit and photometric analyses. An Assumed Circular Orbit formulation provided considerable refinements in semimajor axis and eccentricity, up to one order of magnitude, when compared to a Gauss Extended method. Regarding the use of initial orbits for future pass predictions, the Assumed Circular Orbit angular positions indicated improvements up to 97.4% in accuracy and 65.7% in consistency over Gauss Extended. A photometric study placed the brightest observed visual magnitude at 3.60 mag, and the faintest visible at 9.47 mag. By converting brightness to a physical size, detected objects were approximately 23.8 meters at the largest and 40.6 centimeters at the smallest. Angles and brightness measurements of LEO satellites with mobile platforms may thus benefit the SDA effort.

Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003. / Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.

A new, novel numerical optimization algorithm is developed, tested, and used to solve difficult numerical problems from the field of astrodynamics. First, a brief review of optimization theory is presented and common numerical optimization techniques are discussed. Then, the new method, called the Learning Approach to Sampling Optimization (LA) is presented. Simple, illustrative examples are given to further emphasize the simplicity and accuracy of the LA method. Benchmark functions in lower dimensions are studied and the LA is compared, in terms of performance, to widely used methods. Three classes of problems from astrodynamics are then solved. First, the N - impulse orbit transfer and rendezvous problems are solved by using the LA optimization technique along with derived bounds that make the problem computationally feasible. This marriage between analytical and numerical methods allows an answer to be found for an order of magnitude greater number of impulses than are currently published. Next, the N -impulse work is applied to design periodic close encounters (PCE) in space. The encounters are defined as an open rendezvous, meaning that two spacecraft must be at the same position at the same time, but their velocities are not necessarily equal. The PCE work is extended to include N -impulses and other constraints, and new examples are given. Finally, a trajectory optimization problem is solved using the LA algorithm and comparing performance with other methods based on two models-with varying complexity-of the Cassini-Huygens mission to Saturn. The results show that the LA consistently outperforms commonly used numerical optimization algorithms.

No description available.

Cubesats are picosatellites that have a mass of less than 1.3kg and have a shape of acube. As a result of their low cost of development and launch, cubesats are gainingpopularity in industry and academia. These satellites are also a cost-efective way forspace technology demonstrations. University of Würzburg has a longstanding cubesatprogram started with the launch of UWE-1 in 2005. This was followed by UWE-2 andUWE-3. Several technologies were tested and validated using the UWE platform. Thelast mission UWE-3 has successfully tested an attitude control system.In the next mission, UWE-4 will demonstrate an orbit control system. Being a picosatellite as small as this one (10 x 10 x 10cm 3 and 1kg) brings new challenges intodi↵erent aspects of satellite design, development, control and operation. The orbit con-trol of such a satellite is one of the problems that should be tackled. Being such a smallsatellite means having less propellant mass and much smaller thrusters than conventionalsatellites. These should be addressed in the orbit control. UWE-4 will take advantage of four NanoFEEP thrusters, on one side. Because of theiraccuracy and functionality, these thrusters can be used to implement a continuous thrustsystem. They are also a good choice because of their low energy usage. This work startswith the preparation that was needed to implement a control system. Then explains thestate of the art for continuous thrust control systems. Implements two di↵erent methods,based on perfect control and discusses the outcome. It discuses the limiting factors, likefuel mass, available electrical energy and their e↵ect on the controller performance andconcludes with recommendation for the future researches. / UWE-4

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 $^{\circ}$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.

In recent times, as the National Aeronautics and Space Administration (NASA) focuses on establishing a sustained presence in cislunar space, there has been an increase in planned missions to the cislunar vicinity for lunar exploration. Due to this increase in planned missions, the use of cislunar structures available in the Circular Restricted Three-Body Problem (CRTBP) has become of greater interest. Traditionally, transfers that leverage CRTBP structures in the cislunar vicinity have been generated as point designs. As a consequence of the non-linearity of this model, transitioning these point designs to other epochs or mission scenarios is non-trivial. Hence, a trade space of transfer solutions, that leverage the underlying dynamics, is of interest for rapid mission design. In this study, numerical methods and dynamical systems theory are leveraged to extract available dynamical structures in the model, which are subsequently exploited for transfer design. A trade space of relatively low time of flight, two-maneuver transfers, from a 500 km altitude Low Earth Orbit (LEO) to the Earth-Moon L1 Lyapunov orbit family is generated and analyzed.

<div>To meet the challenges posed by future space exploration activities, relative satellite motion techniques and capabilities require development to incorporate dynamically complex regimes. ?Specific relative motion applications including orbital rendezvous and spacecraft loitering will play a significant role in NASA's Gateway and Artemis missions which aim to land the ?first woman and next man on the Moon by 2024. In this investigation, relative motion in the restricted 3-body problem is formulated, validated, and tested in a rotating local-vertical-local-horizontal (LVLH)</div><div>frame situated at a target spacecraft and followed by a chaser. Two formulations of the restricted 3-body problem are considered, namely the Circular Restricted 3-Body</div><div>Problem (CR3BP) and the Elliptical Restricted 3-Body Problem (ER3BP). Comparisons between the relative dynamical models in the CR3BP and ER3BP, respectively,</div><div>and other standard relative motion sets of equations such as the Euler-Hill (HCW) model and the Linear Equations of Relative Motion (LERM) are accomplished to identify limitations and inaccuracies pertaining to the in orbits that exist in the CR3BP and ER3BP, respectively. Additionally, the relative motion equations are linearized to develop computational tools for solutions to the rendezvous and space loitering problems in the Earth-Moon system.</div>

This thesis is a feasibility study for a crewed mission to a Near Earth Asteroid (NEA). An alternate dual-launch architecture is proposed and analyzed against a more established architecture. Instead of a rendezvous in a low-Earth parking orbit, the new architecture performs the rendezvous while the two spacecraft are on an Earth-escape trajectory to the destination NEA. After selecting a target asteroid, 2000 SG344, each architecture will have its best mission compared to the best mission of the other architecture. Using the new architecture, a mission is created to the chosen NEA, 2000 SG344. A back-up Orion MPCV and a Habitation Module are launched first on a cargo configuration SLS. A crew of two astronauts is launched two hours later in the primary Orion MPCV by a crewed configuration SLS. Both of these launches are on an Earth-escape trajectory and begin rendezvous after two full days in outer space. The completed spacecraft journeys the rest of the trip to the NEA. For a period of eight days, the spacecraft remains in a tight control sphere near the asteroid by using a control algorithm and the rendezvous thrusters. The astronauts have this period to perform their EVAs and accomplish their mission objectives at the NEA. The spacecraft then departs the NEA and returns to Earth. The entire mission is 134 days and requires 2.054 km/s of Delta-v maneuvers to complete. An analysis of multiple Lambert's methods is also done due to their extensive use in this thesis. Many of the most popular Lambert algorithms are compared by evaluating each on its accuracy, speed, and singularities. The best Lambert method to use for the orbital analysis in this paper is Battin's method because it is accurate, quick, and robust for all cases that will be observed.

This work presents the first measurement of turbulent burning velocities of a highly-turbulent compressible standing flame induced by shock-driven turbulence in a Turbulent Shock Tube. High-speed schlieren, chemiluminescence, PIV, and dynamic pressure measurements are made to quantify flame-turbulence interaction for high levels of turbulence at elevated temperatures and pressure. Distributions of turbulent velocities, vorticity and turbulent strain are provided for regions ahead and behind the standing flame. The turbulent flame speed is directly measured for the high-Mach standing turbulent flame. From measurements of the flame turbulent speed and turbulent Mach number, transition into a non-linear compressibility regime at turbulent Mach numbers above 0.4 is confirmed, and a possible mechanism for flame generated turbulence and deflagration-to-detonation transition is established.