Although lightning-generated whistlers have been studied for nearly a century, there are still questions to be answered. It is clear that, at least in a certain frequency range, these waves significantly contribute to the overall wave intensity in the inner magnetosphere. They also influence distribution functions of energetic particles in the Van Allen radiation belts. Thanks to the on-board implemented neural network for automated whistler detection, we can relate electromagnetic wave and energetic particle flux measurements performed by the low-altitude DEMETER spacecraft with the number and dispersion of whistlers detected during a certain time interval. We distinguish the cases with high and low whistler occurrence and we use this information to determine the overall effect of lightning-generated whistlers. Powered by TCPDF (www.tcpdf.org)

Chemical thrusters are widely used in spacecraft for attitude control and orbital manoeuvres. They produce a plume of neutral gas which produces ions via photoionisation and charge exchange. Measurements of local plasma properties will be aff ected by perturbations caused by the coupling between the newborn ions and the plasma. A model of neutral expansion has been used in conjunction with a fully three-dimensional hybrid code to study the evolution and ionisation over time of the neutral cloud produced by the ring of a mono-propellant hydrazine thruster as well as the interactions of the resulting ion cloud with the ambient solar wind. A parameter survey was performed for varying angles of injection and injection rates, particle kinetics were also investigated. Results are presented which show that the plasma in the region near to the spacecraft will be perturbed for an extended period of time with the formation of an interaction region around the spacecraft, a moderate amplitude density bow wave bounding the interaction region and evidence of an instability at the forefront of the interaction region which causes clumps of ions to be ejected from the main ion cloud quasi periodically and the ways in which these features are modifi ed by the degree of solar wind mass loading and the relative orientation of the magnetic fi eld to the angle of injection. This may a ffect Solar Probe Plus for a signifi cant duration as data taking and delicate sensory equipment may be required to cease operation until local fluctuations return to a more moderate level. The scale of the fluctuations seen are dependent upon the duration of the thruster ring and the speci fic geometry and therefore e ffects may vary in-situ.

"A study is presented detailing the simulation of a drag-free follow-on mission to NASA’s Gravity Recovery and Climate Experiment (GRACE). This work evaluates controller performance, as well as thrust, power, and propellant mass requirements for drag-free spacecraft operation at orbital altitudes of 160 - 225 kilometers. In addition, sensitivities to thermospheric wind, GPS signal accuracy and availability of ephemeris data are studied. Orbital dynamics were modeled in Matlab and take into account 2 body gravity effects, J2-J6 non-spherical Earth effects, atmospheric drag and control thrust. A drag model is used in which the drag acceleration is a function of the spacecraft’s relative velocity to the atmosphere, and a “drag parameter,” which includes the spacecraft’s drag coefficient and local mass density of the atmosphere. A MSISE-90 atmospheric model is used to provide local mass densities as well as free stream flow conditions for a Direct Simulation Monte Carlo drag analysis used to validate the spacecraft drag coefficient. The controller is designed around an onboard inertial sensor which uses a freely floating reference mass to measure deviations in the spacecraft position, resulting from non-gravitational forces, from a desired target orbit. Thruster (control actuator) models are based on two different Hall thrusters for providing the orbital along-track acceleration, colloid thrusters for the normal acceleration, and a miniature xenon ion thruster (MiXI) for the cross-track acceleration. The most demanding propulsion requirements correspond to the lowest altitude considered, 160 kilometers. At this altitude the maximum along-track thrust component is calculated to be 98 millinewtons with a required dynamic (throttling) response of 41 mN/s. The maximum position error at this altitude was shown to be in the along-track direction with a magnitude of 3314.9 nanometers and a peak spectral content of 1800 nm/sqrt(Hz) at about 0.1 Hz. At 225 kilometers, the maximum along-track thrust component reduces to 10.3 millinewtons. The maximum dynamic response at this altitude is 4.23 mN/s. The maximum along-track position error is reduced to 367.9 nanometers with a spectral content peak of 40 nm/sqrt(Hz) at 0.1 Hz. For all altitudes, the maximum state errors increase as the mission length increases, however, higher altitude missions show less of a maximum displacement error increase over time than those of lower orbits. The ability of a colloid thruster to control the normal drift is found to be dependent on how frequently the spacecraft state data is updated. Reducing the period between updates from 10 seconds to 1 second reduces the maximum normal state error component from 199 nanometers to less than 32 nanometers, suggesting that spacecraft state update frequency could be a major driver in keeping the spacecraft on the target trajectory. Sensitivity of maximum required thrust and accumulated sensor error to measurement uncertainty is found to be less of a driver than state update frequency. A ‘worst case” thermospheric wind gust was modeled to show the increase on propulsion requirements if such an event were to occur. At 200 kilometers, maximum winds have been measured to be in increase of 650 m/s in the westward direction in the southern pole region. Assuming the majority of the 650 m/s gust occurs over a 4 second time span, the maximum required cross-track thrust at 200 kilometers increases from 1.12 to 2.01 millinewtons. This large increase may drive the thruster choice for a drag-free mission at a similar altitude. For the spacecraft point design considered with a propellant mass fraction of 0.18, the mission lifetime for the 160 km case was calculated to be 0.76 years. This increases 2.27 years at an altitude of 225 km."

In space and astrophysical plasmas, turbulence is responsible for transferring energy from large scales driven by violent events or instabilities, to smaller scales where turbulent energy is ultimately converted into plasma heat by dissipative mechanisms. In the inertial range, the self-similar turbulent energy cascade to smaller spatial scales is driven by the nonlinear interaction between counterpropagating Alfvén waves, denoted Alfvén wave collisions. For the more realistic case of the collision between two initially separated Alfvén wavepackets (rather than previous idealized, periodic cases), we use a nonlinear gyrokinetic simulation code, AstroGK, to demonstrate three key properties of strong Alfvén wave collisions: they (i) facilitate the perpendicular cascade of energy and (ii) generate current sheets self-consistently, and (iii) the modes mediating the nonlinear interaction are simply Alfvén waves. Once the turbulent cascade reaches the ion gyroradius scale, the Alfvén waves become dispersive and the turbulent energy starts to dissipate, energizing the particles via wave-particle interactions with eventual dissipation into plasma heat. The novel Field-Particle Correlation technique determines how turbulent energy dissipates into plasma heat by identifying which particles in velocity-space experience a net gain of energy. By utilizing knowledge of discrete particle arrival times, we devise a new algorithm called PATCH (Particle Arrival Time Correlation for Heliophysics) for implementing a field-particle correlator onboard spacecraft. Using AstroGK, we create synthetic spacecraft data mapped to realistic phase-space resolutions of modern spacecraft instruments. We then utilize Poisson statistics to determine the threshold number of particle counts needed to resolve the velocity-space signature of ion Landau damping using the PATCH algorithm.

Spacecraft in orbit experience temperature swings close to 240 K as the craft passes from the shadow of the Earth into direct sunlight. To regulate the craft’s internal energy, large radiators eject unwanted energy into space using radiation transfer. The amount of radiation emitted is directly related to the topology of the radiator design. Deformable structures such as those made with origami tessellation patterns offer a mechanism to control the quantity of energy being emitted by varying the radiator shape. Three such patterns, the Waterbomb, Huffman Waterbomb, and Huffman Stars-Triangles, can be folded into tubes. Origami tubes offer greater control and simplicity of design than flat radiators. Using FLUENT, Origami Simulator, and Solidworks to first simulate and then analyze the flow of a thermal fluid through the patterns and the radiation emitted from the created bodies, it was determined that the Waterbomb pattern achieved a 17.6 percent difference in emitted radiation, over a 2 percent change in fold. The Huffman Waterbomb pattern displayed a 42.7 percent difference in emitted radiation over a 20 percent change of fold. The simulations demonstrated both the feasibility and benefits of the origami designed tubes.

<p>Spacecraft on interplanetary missions or advanced satellites orbiting the Earth all require propulsion systems to complete their missions. Introducing microelectromechanical systems technology to the space industry will not only reduce size and weight of the propulsion system, but can also increase the performance of the mission.</p><p>Fluid handling systems are used in chemical and electric propulsion. Some components incorporated in a fluidic handling system are presented and evaluated in this work.</p><p>Microsystems are very sensitive to contamination. Reliable, robust, and easily integrated filters were modeled, manufactured, and experimentally verified.</p><p>A fluid connector, designed to withstand large temperature variations and aggressive propellants was manufactured and characterized. Similar designs was also be used as a thermally activated minute valve.</p><p>The feasibility of a cold gas system for precise attitude control has been demonstrated. Steps towards improving the performance (from specific im-pulse 45 s) have been taken, by the integration of suspended heater elements.</p><p>For electric propulsion, two thermally regulated flow restrictors have been characterized. These devices can fine-tune the propellant flow to e.g. an ion engine.</p><p>A single-use valve using a soldered seal has also been successfully dem-onstrated within a pressure range of 5 to 100 bar.</p><p>The microsystem-based propulsion systems of tomorrow’s spacecraft need to be demonstrated in space, in order to gain necessary credibility. </p>

A perfect duality exists between the problem of space-based orbit determination from line-of-sight measurements and the problem of designing an interplanetary autonomous navigation system. Mathematically, these two problems are equivalent. Any method solving the first problem can be used to solve the second one and, vice versa. While the first problem estimates the observed unknown object orbit using the known observer orbit, the second problem does exactly the opposite (e.g. the spacecraft observes a known visible planet). However, in an interplanetary navigation problem, in addition to the measurement noise, the following "perturbations" must be considered: 1) light-time effect due to the finite speed of light and large distances between the observer and planets, and 2) light aberration including special relativistic effect. These two effects require corrections of the initial orbit estimation problems. Because of the duality problem of space-based orbit determination, several new techniques of angles-only Initial Orbit Determination (IOD) are here developed which are capable of using multiple observations and provide higher orbit estimation accuracy and also they are not suffering from some of the limitations associated with the classical and some newly developed methods of initial orbit determination. Using multiple observations make these techniques suitable for the coplanar orbit determination problems which are the case for the spacecraft navigation using visible planets as the solar system planets are all almost coplanar. Four new IOD techniques were developed and Laplace method was modified. For the autonomous navigation purpose, Extended Kalman Filter (EKF) is employed. The output of the IOD algorithm is then used as the initial condition to extended Kalman filter. The two "perturbations" caused by light-time effect and stellar aberration including special relativistic effect also need to be taken into consideration and corrections should be implemented into the extended Kalman filter scheme for the autonomous spacecraft navigation problem.

The Generic Nanosatellite Bus (GNB) is a spacecraft platform designed to accommodate the integration of diverse payloads in a common housing of supporting components. The development of the GNB at the Space Flight Laboratory (SFL) under the Canadian Advanced Nanospace eXperiment (CanX) program provides accelerated access to space while reducing non-recurring engineering (NRE) costs. The work presented herein details the development of the attitude determination and control subsystem (ADCS) of the GNB. Specific work on magnetorquer coil assembly, integration, and testing (AIT) and reaction wheel testing is included. The embedded software development and unit-level testing of the GNB sun sensors are discussed. The characterization of the AeroAstro star tracker is also a major focus, with procedures and results presented here. Hardware models were developed and incorporated into SFL's in-house high-fidelity attitude dynamics and control simulation environment. This work focuses on specific contributions to the CanX-3, CanX-4&5, and AISSat-1 nanosatellite missions.

This thesis presents a method for managing periods of communication loss between multiple spacecraft in formation flying (MSFF), and analyzes the effects of this method on the stability of the formation keeping control algorithm. The controller of interest in this work in an adaptive nonlinear controller, where synchronization is also incorporated to force the position tracking errors to converge to zero at the same rate. The communication loss compensation technique proposed in this thesis is to use the previously communicated data in lieu of the lost data, which is an effective and computationally-efficient technique that is advantageous for small satellites. The performance parameter of interest in this research is the maximum rate of communication loss that an MSFF system can withstand before going unstable, and this is analyzed theoretically and through simulations. Finally, experiments involving multiple robots in formation with communication loss are conducted, and the results are presented.

The Generic Nanosatellite Bus (GNB) is a spacecraft platform designed to accommodate the integration of diverse payloads in a common housing of supporting components. The development of the GNB at the Space Flight Laboratory (SFL) under the Canadian Advanced Nanospace eXperiment (CanX) program provides accelerated access to space while reducing non-recurring engineering (NRE) costs. The work presented herein details the development of the attitude determination and control subsystem (ADCS) of the GNB. Specific work on magnetorquer coil assembly, integration, and testing (AIT) and reaction wheel testing is included. The embedded software development and unit-level testing of the GNB sun sensors are discussed. The characterization of the AeroAstro star tracker is also a major focus, with procedures and results presented here. Hardware models were developed and incorporated into SFL's in-house high-fidelity attitude dynamics and control simulation environment. This work focuses on specific contributions to the CanX-3, CanX-4&5, and AISSat-1 nanosatellite missions.