Sheath ionization model of beam emissions from large spacecraft

An analytical model of the charging of a spacecraft emitting electron and ion beams has been applied to the case of large spacecraft. In this model, ionization occurs in the sheath due to the return current. Charge neutralization of spherical space charge flow is examined by solving analytical equations numerically. Parametric studies of potential large spacecraft are performed. As in the case of small spacecraft, the ions created in the sheath by the returning current play a large role in determining spacecraft potential

Active spacecraft potential control system selection for the Jupiter orbiter with probe mission

It is shown that the high flux of energetic plasma electrons and the reduced photoemission rate in the Jovian environment can result in the spacecraft developing a large negative potential. The effects of the electric fields produced by this charging phenomenon are discussed in terms of spacecraft integrity as well as charged particle and fields measurements. The primary area of concern is shown to be the interaction of the electric fields with the measuring devices on the spacecraft. The need for controlling the potential of the spacecraft is identified, and a system capable of active control of the spacecraft potential in the Jupiter environment is proposed. The desirability of using this system to vary the spacecraft potential relative to the ambient plasma potential is also discussed. Various charged particle release devices are identified as potential candidates for use with the spacecraft potential control system. These devices are evaluated and compared on the basis of system mass, power consumption, and system complexity and reliability

A charging model for three-axis stabilized spacecraft

A charging model was developed for geosynchronous, three-axis stabilized spacecraft when under the influence of a geomagnetic substorm. The differential charging potentials between the thermally coated or blanketed outer surfaces and metallic structure of a spacecraft were determined when the spacecraft was immersed in a dense plasma cloud of energetic particles. The spacecraft-to-environment interaction was determined by representing the charged particle environment by equivalent current source forcing functions and by representing the spacecraft by its electrically equivalent circuit with respect to the plasma charging phenomenon. The charging model included a sun/earth/spacecraft orbit model that simulated the sum illumination conditions of the spacecraft outer surfaces throughout the orbital flight on a diurnal as well as a seasonal basis. Transient and steady-state numerical results for a three-axis stabilized spacecraft are presented

Method and apparatus for neutralizing potentials induced on spacecraft surfaces

A potential induced on the surface of an orbiting spacecraft is neutralized to the potential of a plasma through which the spacecraft is traveling by directing charged particles into the plasma from the spacecraft surface. The induced potential occurs in response to bombardment of the spacecraft surface by ambient charged particles which may be negative or positive. The charged particles directed into the plasma from the surface have the same polarity as the induced potential to provide the neutralization. The invention can be utilized to maintain different, electrically isolated segments of a spacecraft surface at the same potential to prevent electric discharges between the different parts and thereby protect electric circuits within the spacecraft

Deflecting small asteroids using laser ablation : Deep space navigation and asteroid orbit control for LightTouch2 Mission

This paper presents a low-cost, low mass, mission design to successfully intercept and deflect a small and faint, 4 m in diameter asteroid. Intended to be launched after 2025, the laser-ablating mission, LightTouch2 will be used to deflect the orbit of the asteroid by at least 1 m/s. This will be achieved with a total mission lifetime of less than three years. Analysis includes the initial approach of the spacecraft, the operations of the laser at an optimal spacecraft-to-asteroid distance of 50 m and the relative orbit of the spacecraft that flies in formation with the asteroid. Analysis includes line-of-sight measurements with radiometric tracking from ground station to improve the trajectory estimate and observability of the spacecraft, collision avoidance and mapping strategies. The spacecraft will also need optimal discrete control. This is achieved by impulse-bit manoeuvres used to account for the perturbations caused by the resultant thrust on the asteroid, plume impingement, laser recoil and solar radiation pressure. The spacecraft controls its trajectory within a 1 m box from the reference trajectory to enable the laser to optimally focussing the laser beam. The proposed approach uses an unscented Kalman filter to estimate the relative spacecraft-asteroid position, velocity and perturbative acceleration

Attitude stabilization of a rigid spacecraft using two momentum wheel actuators

It is well known that three momentum wheel actuators can be used to control the attitude of a rigid spacecraft and that arbitrary reorientation maneuvers of the spacecraft can be accomplished using smooth feedback. If failure of one of the momentum wheel actuators occurs, it is demonstrated that two momentum wheel actuators can be used to control the attitude of a rigid spacecraft and that arbitrary reorientation maneuvers of the spacecraft can be accomplished. Although the complete spacecraft equations are not controllable, the spacecraft equations are small time locally controllable in a reduced nonlinear sense. The reduced spacecraft dynamics cannot be asymptotically stabilized to any equilibrium attitude using a time-variant continuous feedback control law, but discontinuous feedback control strategies are constructed which stabilize any equilibrium attitude of the spacecraft in finite time. Consequently, reorientation of the spacecraft can be accomplished using discontinuous feedback control

Nuclear radiation problems, unmanned thermionic reactor ion propulsion spacecraft

A nuclear thermionic reactor as the electric power source for an electric propulsion spacecraft introduces a nuclear radiation environment that affects the spacecraft configuration, the use and location of electrical insulators and the science experiments. The spacecraft is conceptually configured to minimize the nuclear shield weight by: (1) a large length to diameter spacecraft; (2) eliminating piping penetrations through the shield; and (3) using the mercury propellant as gamma shield. Since the alumina material is damaged by the high nuclear radiation environment in the reactor it is desirable to locate the alumina insulator outside the reflector or develop a more radiation resistant insulator

Precise Model for Small-Body Thermal Radiation Pressure Acting on Spacecraft

A precise representation of small-body surface thermal radiation pressure effects acting on orbiting spacecraft is discussed. The proposed framework takes advantage of a general Fourier series expansion to compute small-body surface thermal radiation pressure. Fourier series expansion has been used before for the precise representation of solar radiation pressure effects on spacecraft orbiting small bodies. This framework takes into account the geometric relationship of orbiting spacecraft with the small-body surface, surface thermal parameters of the small body, and the shape and surface properties of spacecraft allowing for the computation of thermal radiation pressure, which may also be used for the generation of precise orbit determination solutions. After presenting the general model, an example application of the model for the OSIRIS-REx spacecraft in orbit about Asteroid (101955) Bennu is provided. Simulation studies were used to evaluate the effect of mismodeling of thermal radiation pressure on the spacecraft and study the use of the proposed method for generating precise orbit determination solutions