Spacecraft Engineers Research Articles

Automated Plume Sentry Observations During International Space Station Thermal Control System Venting

Contamination from outgassed materials, venting, leaks, and impinging thruster plumes can cause deleterious effects to sensitive experiments on the International Space Station. In low Earth orbit, intermolecular collisions between neutral contaminants and ions in the ambient ionospheric environment result in an elevation in the local charged particle density through a process called charge exchange (CEX). Of particular concern to spacecraft engineers is the acceleration of these ionized neutrals to sensitive negatively charged spacecraft surfaces. The increase in plasma density due to CEX can be readily measured using plasma diagnostics such as electrostatic energy analyzers. These instruments are capable of monitoring charged particle flux and are sensitive enough to measure small changes in local density, thus allowing for the detection of CEX contaminants. The automated plume sentry (APS) is one such instrument, and it has successfully detected CEX plasma generated by the planned ammonia venting of the external active thermal control system. The data collected by the APS will support the development of improved computational models to predict the propagation of outgassed materials and plumes in the ionosphere and complex gas–surface interactions resulting in spacecraft surface contamination.

Thermal response and surface recession of a carbon-phenolic charring heatshield of spacecraft: Numerical simulation and validation

Aero-thermal heating experienced by spacecraft as they enter planetary atmosphere is exorbitant. The response of a sacrificial thermal shielding system of a spacecraft to extreme heating depends on the unsteady dynamics of thermo-physico-chemical interaction of shielding material with high enthalpy plasma. This paper addresses a transient analysis being performed to elucidate these interaction phenomena and to approximate the thermal response and the surface recession of an iso-q model made of low-density, high-porosity carbon/phenolic ablative material. The newly developed finite element material response code of this study solved time-dependent governing equations, including two energy conservation equations (for internal- and surface energy). A three-component chemical decomposition model approximated the material decomposition rate. Darcy's law predicted the direction of pyrolysis gas flow as well as the pressure gradient across the homogeneously porous solid material. A shaver re-meshing technique refined the ablation front posterior to the surface recession. Benchmark solutions were computed and compared against available solutions to interrogate the fidelity of the proposed code. For an identical surface energy boundary condition, the proposed code predicted a material temperature in excellent agreement with the thermal response data inferred from the material response code “Amaryllis”. Besides, the predicted char contours were similar to that of a high enthalpy air-plasma-exposed iso-q specimen. The outcomes of this study will reinforce the confidence of spacecraft engineers on the proposed code for the aero-thermal analysis of next generation sacrificial thermal shielding systems.

Theory of Power Generation From Spacecraft Charging

Space plasmas, though tenuous, can lead to the development of large magnitude potentials on spacecraft surfaces relative to the ambient plasma in a process known as spacecraft charging. This effect poses a threat to spacecraft when surfaces charge differentially as it increases the risk of electrical arcing. Spacecraft engineers have primarily focused on the mitigation of charging and not its exploitation. The potentials and current collected by the spacecraft present an opportunity to harvest power from the space environment. Although the current density and floating potential due to incident charged species in the orbit-motion-limited model are identical for all spacecraft surfaces with similar geometries, material property differences between surfaces or charge control devices break the symmetry and allow for net current to flow. Differences in the secondary electron yield, in particular, produce strong differences in floating potential even when the spacecraft is eclipsed or far from the sun. This article presents a theory for computing the power available to harvest from spacecraft charging given the parameters of ambient space plasmas and surface material properties. Given a load connecting two differentially charged surfaces, optimal power is harvested by matching the load resistance to the input plasma impedance. While this holds for any model of the ambient plasma, the orbit-motion-limited theory yields closed-form solutions to the charging equations for spherical and planar surfaces. These solutions provide an upper bound on generated power that scales with the density and temperature of the cold electron population. An environment-specific optimal anode-to-cathode area ratio exists that minimizes the input impedance to the harvesting circuit. Deployable anodes are recommended to satisfy this condition and increase the amount of power harvested. Energy harvesting from spacecraft charging can be used in the Jovian magnetosphere to generate on the order of 10 mW/m <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">2</sup> of power.

Testing satellite control systems with drones

In space, it is usually not possible to repair equipment when things go wrong. While astronauts ventured into space aboard the Space Shuttle to repair the infamously flawed Hubble Space Telescope, these types of repair missions are extremely expensive and rare. Limited repair options make spacecraft engineers turn their attention to reliability in an attempt to avoid failures altogether. Building a reliable system starts with a robust design, incorporates proven parts, and—most importantly—involves a lot of testing.

Formal Verification of Spacecraft Control Programs

Verification of correctness of control programs is an essential task in the development of space electronics; it is difficult and typically outweighs design and programming tasks in terms of development hours. This article presents a verification approach designed to help spacecraft engineers reduce the effort required for formal verification of low-level control programs executed on custom hardware. The verification approach is demonstrated on an industrial case study. We present a REDuced instruction set for Fixed-point and INteger arithmetic (REDFIN), a processing core used in space missions, and its formal semantics expressed using the proposed metalanguage for state transformers, followed by examples of verification of simple control programs.

Design of an ASIC Chip for Spacecraft Data System

ASIC (Application Specific Integrated Circuit) technology is mature and widely used in many fields. It is a trend to integrate some common spacecraft engineering requirements into one chip to take the place of FPGA and some digital chips, which can miniaturize spacecraft avionics and reduce the repeated work for spacecraft engineers. Based on the analysis on small satellites application requirements, the system design of an ASIC chip for spacecraft data system is proposed. Firstly, the chip system architecture is described. Secondly, four key technologies of the ASIC chip design are presented in detail. They are the design of the chip's operating modes, the design of IP cores, the design of reliability, and the design of low power consumption. Generality, adaptability and independent intellectual property make this ASIC different and special from other space chips. Four operating modes are optional: 1553B bus control mode, PCI control mode, ISA control mode, internal CPU control mode. Besides, there are 1553B bus protocol IP cores, CAN bus interface, and abundant peripherals, which can satisfy most of the routine engineering requirements in small satellites. Finally, the chips status and some promising applications are described.

Planetary seismology—Expectations for lander and wind noise with application to Venus

The amplitudes of seismic signals on a planetary surface are discussed in the context of observable physical quantities – displacement, velocity and acceleration – in order to assess the number of events that a sensor with a given detection threshold may capture in a given period. Spacecraft engineers are generally unfamiliar with expected quantities or the language used to describe them, and seismologists are rarely presented with the challenges of accommodation of instrumentation on spacecraft. This paper attempts to bridge this gap, so that the feasibility of attaining seismology objectives on future missions – and in particular, a long-lived Venus lander – can be rationally assessed.For seismometers on planetary landers, the background noise due to wind or lander systems is likely to be a stronger limitation on the effective detection threshold than is the instrument sensitivity itself, and terrestrial data on vehicle noise is assessed in this context. We apply these considerations to investigate scenarios for a long-lived Venus lander mission, which may require a mechanical cooler powered by a Stirling generator. We also consider wind noise: the case for decoupling of a seismometer from a lander is strong on bodies with atmospheres, as is the case for shielding the instrument from wind loads. However, since the atmosphere acts on the elastic ground as well as directly on instruments, the case for deep burial is not strong, but it is important that windspeed and pressure be documented by adequate meteorology measurements.

New radiation environment and effects models in the European Space Agency's Space Environment Information System (SPENVIS)

The European Space Agency (ESA) Space Environment Information System (SPENVIS) provides standardized access to models of the hazardous space environment through a user‐friendly Web interface, available at http://www.spenvis.oma.be/. SPENVIS is designed to help spacecraft engineers perform rapid analyses of environmental problems and, with extensive documentation and tutorial information, allows engineers with relatively little familiarity with the models to produce reliable results. SPENVIS is based on internationally recognized standard models and methods in many domains. It uses an ESA‐developed orbit generator to produce orbital point files necessary for many different types of problem. It has various reporting and graphical utilities and extensive help facilities. SPENVIS also contains an active, integrated version of the European Cooperation for Space Standardization Space Environment Standard, and access to in‐flight data. Apart from radiation and plasma environments, SPENVIS includes meteoroid and debris models, atmospheric models (including atomic oxygen), and magnetic field models. Recently, the radiation sources and effects module have been rewritten, providing enhanced mission analysis features, new radiation effects models, and streamlined navigation between various model pages.

Space Robotics

In this first of three short papers, I introduce some of the basic concepts of space engineering with an emphasis on some specific challenging areas of research that are peculiar to the application of robotics to space development and exploration. The style of these short papers is pedagogical and this paper stresses the unique constraints that space application imposes. This first paper is thus a general introduction to the nature of spacecraft engineering and its application to robotic spacecraft. I consider the constraints and metrics used by spacecraft engineers in the design of spacecraft and how these constraints impose challenges to the roboticist. The following two papers consider specific robotics issues in more detail.

Radiation belt modelling in the framework of space weather effects and forecasting

The Earth's trapped radiation belts were discovered at the beginning of the space age and were immediately recognised as a considerable hazard to space missions. Consequently, considerable effort was invested in building models of the trapped proton and electron populations, culminating in the NASA AP-8 and AE-8 models which have been de facto standards since the 1970s. The CRRES mission has demonstrated that the trapped radiation environment is much more complex than the static environment described by the old models. Spatial and especially temporal variations were shown to be much more important than previously thought, and to require more complex models than those in use at that time. Such models are now becoming available, but they are as yet limited in spatial and temporal coverage. It is vital to coordinate future modelling efforts in order to develop new standard models. The lack of standardisation of radiation belt models complicates their use in engineering applications where particle fluxes are needed as input to radiation effects models. ESA's SPace ENVironment Information System (SPENVIS) provides standardised access to models of the hazardous space environment, including but not limited to radiation effects, through a user-friendly World Wide Web interface. The interface includes parameter input with extensive defaulting, definition of user environments, streamlined production of results (both in graphical and textual form), background information, and on-line help. The system can be accessed at the WWW site http://www.spenvis.oma.be/spenvis/. SPENVIS Is designed to help spacecraft engineers perform rapid analyses of environmental problems and, with extensive documentation and tutorial information, allows engineers with relatively little familiarity to produce reliable results. It has been developed in response to the increasing pressure for rapid-response tools for system engineering, especially in low-cost commercial and educational programmes.

Towards a new model of the interplanetary meteoroid environment

Improved models of the interplanetary meteoroid environment enjoy the interest of both spacecraft engineers and dust researchers. The engineers need it for risk assessments for their spacecraft instruments. Modelling dynamical and collisional evolution of interplanetary dust should lead to a match with observations, and an empirical model can be a good mediator between physical models and sparse observational data. Our current effort is directed towards the construction of a new model of the interplanetary meteoroid environment based on a number of observational data sets including in-situ dust flux measurements onboard spacecraft, radar meteor surveys and thermal emission of zodiacal dust. In contrast to earlier models, we use long-term particle dynamics to define populations for the new model. Based on these populations, we have constructed a prototype model which reasonably fits in-situ impact counts by Galileo and Ulysses dust experiments.

Euler, Jacobi, and missions to comets and asteroids

Whenever a freely spinning body is found in a complex rotational state, this means that either the body experienced some interaction within its relaxation-time span, or that it was recently “prepared” in a non-principal state. Both options are encountered in astronomy where a wobbling rotator is either a recent victim of an impact or a tidal interaction, or is a fragment of a disrupted progenitor. Another factor (relevant for comets) is outgassing. By now, the optical and radar observational programmes have disclosed that complex rotation is hardly a rare phenomenon among the small bodies. Due to impacts, tidal forces and outgassing, the asteroidal and cometary precession must be a generic phenomenon: while some rotators are in the state of visible tumbling, a much larger amount of objects must be performing narrow-cone precession not so easily observable from the Earth. The internal dissipation in a freely precessing top leads to relaxation (gradual damping of the precession) and sometimes to spontaneous changes in the rotation axis. Recently developed theory of dissipative precession of a rigid body reveals that this is a highly nonlinear process: while the body is precessing at an angular rate ω, the precession-caused stresses and strains in the body contain components oscillating at other frequencies. Dependent upon the spin state, those frequencies may be higher or, most remarkably, lower than the precession rate. In many states dissipation at the harmonics is comparable to or even exceeds that at the principal frequency. For this and other reasons, in many spin states the damping of asteroidal and cometary wobble happens faster, by several orders, than believed previously. This makes it possible to measure the precession-damping rate. The narrowing of the precession cone through the period of about a year can be registered by the currently available spacecraft-based observational means. We propose an appropriate observational scheme that could be accomplished by comet and asteroid-aimed missions. Improved understanding of damping of excited rotation will directly enhance understanding of the current distribution of small-body spin states. It also will constrain the structure and composition of excited rotators. However, in the near-separatrix spin states a precessing rotator can considerably slow down its relaxation. This lingering effect is similar to the one discovered in 1968 by Russian spacecraft engineers who studied free wobble of a tank with viscous fuel.

Analytical Criterion for Chaotic Dynamics in Flexible Satellites with Nonlinear Controller Damping

In this work, we study the attitude dynamics of a single body spacecraft that is perturbed by the motion of a small flexible appendage constrained to undergo only torsional vibration. In particular, we are interested in the chaotic dynamics that can occur for certain sets of the physical parameter values of the spacecraft when energy dissipation acts to drive the body from minor to major axis spin. Energy dissipation, which is present in all spacecraft systems and is the mechanism that drives the minor to major axis transition, is implemented by a quantitative energy sink that is modeled with a nonlinear controller. We obtain an analytical test for chaos in terms of satellite parameters by Melnikov's method. This analytical criterion provides a useful design tool to spacecraft engineers who are concerned with avoiding potentially problematic chaotic dynamics in their systems. In addition, we show that a spacecraft with a control system designed to provide energy dissipation can exhibit chaos because of the inherent flexibility of its components.

Gamma-ray astronomy — Future prospects

With the advent of Sigma/Granat and of the Compton Gamma Ray Observatory, the gamma-ray band has at last been opened up for astronomy. It is starting to be possible to realize some of the potential that observations in this band have for contributing to our understanding of high energy astrophysics. What, though, of prospects beyond the current generation of gamma-ray instruments? It is ironical that in the part of the spectrum where the diffraction limit is in the micro-arcsecond range, angular resolutions available are still comparable with the angular diameter of the moon. Similarly, none of the current generation of space instruments has energy resolution better than about E ΔE ∼ 10 . Yet this is a regime where resolutions of ∼ 1000 are easily achieved in the laboratory and where the theoretical limits are many orders of magnitude better than this. A major problem is sensitivity. The large detector masses necessary to stop high energy gamma-rays attract both high backgrounds from particle interactions and worried looks from spacecraft engineers. Ingenuity, large lift-off masses and long observations are all needed. The prospects for INTEGRAL and beyond are discussed in the context of developments over the last 20 years and the present status.

An overview of electron and ion beam effects in charging and discharging to spacecraft

The charging and discharging of the SCATHA satellite by means of electro- and ion-beam emissions in the geosynchronous environment were investigated. It was found that during electron beam emissions in sunlight the photoelectrons from nearby surfaces isolated from the satellite body tend to return to the body, thus affecting its potential. It was also found that attempts to discharge a spacecraft by means of electron-beam emission can result in differential charging between dielectric surfaces and spacecraft ground. At low ion-beam currents, the level of spacecraft charging increases as the current increases. Beyond a critical current, the level decreases with a nonmonotonic current-voltage behavior. Abundant neutrals are present in the ion beam. Low-energy ions generated by charge exchange return to the spacecraft and reduce the charging level. It was also shown that discharging by means of low-energy plasma-beam emission is more efficient than using an electron or ion beam alone. This conclusion is important for spacecraft engineers designing automatic discharge systems for future spacecraft. >

Expanding the Limits of the Calculator Display

The decimal expansions of 1/7 or 5/13 can be found immediately on the average electronic calculator. These decimal expansions are limited to six digits before they begin to repeat. But what do we say to students who want to explore the repeating properties of rational numbers? What would we answer if they asked for the decimal expansion of 1/17 or 3/23? Do we just say, “Keep dividing and you'll find the answer”? Hardly. Long decimal expansions are usually left to actuaries who don't want to mismanage thousands of dollars over the life of an annuity or mortgage table, or to spacecraft engineers who don't want to miss their targets in outer space.