Table of Contents
Series Page
Title Page
Copyright
Note from the Series Editor
Foreword
Chapter 1: Introduction
References
Chapter 2: Introduction to the Physics of Charging and Discharging
2.1 Physical Concepts
2.2 Electron Environment
2.3 Modeling Spacecraft Charging
2.4 Discharge Characteristics
2.5 Coupling Models
References
Chapter 3: Spacecraft Design Guidelines
3.1 Processes
3.2 Design Guidelines
References
Chapter 4: Spacecraft Test Techniques
4.1 Test Philosophy
4.2 Simulation of Parameters
4.3 General Test Methods
References
Chapter 5: Control and Monitoring Techniques
5.1 Active Spacecraft Charge Control
5.2 Environmental and Event Monitors
References
Chapter 6: Material Notes and Tables
6.1 Dielectric Material List
6.2 Conductor Material List
References
Appendix A: Nomenclature
A.1 Constants and Measurement Units
A.2 Acronyms and Abbreviations
A.3 Defined Terms
A.4 Variables
A.5 Symbols
Appendix B: The Space Environment
B.1 Introduction to Space Environments
B.2 Geosynchronous Environments
B.3 Other Earth Environments
B.4 Other Space Environments
References
Appendix C: Environment, Electron Transport, and Spacecraft Charging Computer Codes
C.1 Environment Codes
C.2 Transport Codes
C.3 Charging Codes
References
Appendix D: Internal Charging Analyses
D.1 The Physics of Dielectric Charging
D.2 Simple Internal Charging Analysis
D.3 Detailed Analysis
D.4 Spacecraft Level Analysis
References
Appendix E: Test Methods
E.1 Electron-Beam Tests
E.2 Dielectric Strength/Breakdown Voltage
E.3 Resistivity–Conductivity Determination
E.4 Simple Volume Resistivity Measurement
E.5 Electron-Beam Resistivity Test Method
E.6 NonContacting Voltmeter Resistivity Test Method
E.7 Dielectric Constant, Time Constant
E.8 Vzap Test [MIL-STD-883G, Method 3015.7 Human Body Model (HBM)]
E.9 Transient Susceptibility Tests
E.10 Component/Assembly Testing
E.11 Surface Charging ESD Test Environments
E.12 System Internal ESD Testing
References
Appendix F: Voyager SEMCAP Analysis
References
Appendix G: Simple Approximations: Spacecraft Surface Charging Equations
References
Appendix H: Derivation of Rule Limiting Open-Circuit Board Area
Reference
Appendix I: Expanded Worst-Case Geosynchronous Earth Environments Descriptions
References
Appendix J: Key Spacecraft Charging Documents
J.1 U.S. Government Documents
J.2 Non-U.S. Government Documents
Color Plates
Index
For further information visit: the book web page http://www.openmodelica.org, the Modelica Association web page http://www.modelica.org, the authors research page http://www.ida.liu.se/labs/pelab/modelica, or home page http://www.ida.liu.se/~petfr/, or email the author at peter.fritzson@liu.se. Certain material from the Modelica Tutorial and the Modelica Language Specification available at http://www.modelica.org has been reproduced in this book with permission from the Modelica Association under the Modelica License 2 Copyright © 1998–2011, Modelica Association, see the license conditions (including the disclaimer of warranty) at http://www.modelica.org/modelica-legal-documents/ModelicaLicense2.html. Licensed by Modelica Association under the Modelica License 2.
Modelica© is a registered trademark of the Modelica Association. MathModelica© is a registered trademark of MathCore Engineering AB. Dymola© is a registered trademark of Dassault Syst`emes. MATLAB© and Simulink© are registered trademarks of MathWorks Inc. Java is a trademark of Sun MicroSystems AB. Mathematica© is a registered trademark of Wolfram Research Inc.
Copyright © 2011 by the Institute of Electrical and Electronics Engineers, Inc.
Published by John Wiley & Sons, Inc., Hoboken, New Jersey. All rights reserved.
Published simultaneously in Canada.
No part of this publication may be reproduced, stored in a retrieval system, or transmitted in any form or by any means, electronic, mechanical, photocopying, recording, scanning, or otherwise, except as permitted under Section 107 or 108 of the 1976 United States Copyright Act, without either the prior written permission of the Publisher, or authorization through payment of the appropriate per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923, (978) 750-8400, fax (978) 750-4744. Requests to the Publisher for permission should be addressed to the Permissions Department, John Wiley & Sons, Inc., 111 River Street, Hoboken, NJ 07030, (201) 748-6011, fax (201) 748-6008, or online at http://www.wiley.com/go/permission.
Limit of Liability/Disclaimer of Warranty: While the publisher and author have used their best efforts in preparing this book, they make no representations or warranties with respect to the accuracy or completeness of the contents of this book and specifically disclaim any implied warranties of merchantability or fitness for a particular purpose. No warranty may be created or extended by sales representatives or written sales materials. The advice and strategies contained herein may not be suitable for your situation. You should consult with a professional where appropriate. Neither the publisher nor author shall be liable for any loss of profit or any other commercial damages, including but not limited to special, incidental, consequential, or other damages.
For general information on our other products and services or for technical support, please contact our Customer Care Department within the United States at (800) 762-2974, outside the United States at (317) 572-3993 or fax (317) 572-4002.
Wiley also publishes its books in a variety of electronic formats. Some content that appears in print may not be available in electronic formats. For more information about Wiley products, visit our web site at www.wiley.com.
Library of Congress Cataloging-in-Publication Data:
Garrett, Henry B.
Guide to mitigating spacecraft charging effects / Henry B. Garrett, Albert C.
Whittlesey. — 1st ed.
p. cm. — (JPL space science and technology series)
Includes bibliographical references and index.
ISBN 978-1-118-18645-9 (hardback)
1. Space vehicles–Electrostatic charging. 2. Electric discharges–Prevention. I. Title.
TL1492.G37 2012
629.47–dc23
2011036330
Note from the Series Editor
The Jet Propulsion Laboratory (JPL) Space Science and Technology Series broadens the range of the ongoing JPL Deep Space Communications and Navigation Series to include disciplines other than communications and navigation in which JPL has made important contributions. The books are authored by scientists and engineers with many years of experience in their respective fields, and lay a foundation for innovation by communicating state-of-the-art knowledge in key technologies. The series also captures fundamental principles and practices developed during decades of space exploration at JPL, and it celebrates the successes achieved. These books will serve to guide a new generation of scientists and engineers.
We would like to thank the Office of the Chief Scientist and Chief Technologist for their encouragement and support. In particular, we would like to acknowledge the support of Thomas A. Prince, former JPL Chief Scientist; Erik K. Antonsson, former JPL Chief Technologist; Daniel J. McCleese, JPL Chief Scientist; and Paul E. Dimotakis, JPL Chief Technologist.
Joseph H. Yuen, Editor-in-Chief
JPL Space Science and Technology Series
Jet Propulsion Laboratory
California Institute of Technology
Foreword
I am very pleased to commend the Jet Propulsion Laboratory (JPL) Space Science and Technology Series, and to congratulate and thank the authors for contributing their time to these publications. It is always difficult for busy scientists and engineers, who face the constant pressures of launch dates and deadlines, to find the time to tell others clearly and in detail how they solved important and difficult problems, so I applaud the authors of this series for the time and care they devoted to documenting their contributions to the adventure of space exploration.
JPL has been NASA's primary center for robotic planetary and deep-space exploration since the Laboratory launched the nation's first satellite, Explorer 1, in 1958. In the 50 years since this first success, JPL has sent spacecraft to all the planets except Pluto, studied our own planet in wavelengths from radar to visible, and observed the universe from radio to cosmic ray frequencies. Current plans call for even more exciting missions over the next decades in all these planetary and astronomical studies, and these future missions must be enabled by advanced technology that will be reported in this series. The JPL Deep Space Communications and Navigation book series captured the fundamentals and accomplishments of these two related disciplines, and we hope that this new series will expand the scope of those earlier publications to include other space science, engineering, and technology fields in which JPL has made important contributions.
I look forward to seeing many important achievements captured in these books.
Charles Elachi, Director
Jet Propulsion Laboratory
California Institute of Technology
Chapter 1
Introduction
This book documents engineering guidelines and design practices that can be used by spacecraft designers to minimize the detrimental effects of spacecraft surface and internal charging in certain space environments. Chapter 2 covers space charging/electrostatic discharge background and orientation; Chapter 3, design guidelines; Chapter 4, spacecraft test techniques; Chapter 5, control and monitoring methods; and Chapter 6, materials that should or should not be considered for charging control. The appendixes contain a collection of useful material intended to support the main body of the document. Despite our desire that this be an all-encompassing guideline, this document cannot do that. It is a narrowly focused snapshot of existing technology, not a research report, and does not include certain related technologies or activities as clarified further below.
In-space charging effects are caused by interactions between the in-flight plasma environment and spacecraft materials and electronic subsystems. Possible detrimental effects of spacecraft charging include disruption of or damage to subsystems (such as power, navigation, communications, or instrumentation) because of field buildup and electrostatic discharge as a result of a spacecraft's passage through the space plasma and high-energy particle environments. Charges can also attract contaminants, affecting thermal properties, optical instruments, and solar arrays; and they can change particle trajectories, thus affecting plasma-measuring instruments. NASA RP-1375, Failures and Anomalies Attributed to Spacecraft Charging (1), lists and describes some spaceflight failures caused by inadequate designs.
This book applies to Earth-orbiting spacecraft that pass through the hazardous regions identified in Figs. 1.1 and 1.2 [medium Earth orbit (MEO), low Earth orbit (LEO), and geosynchronous Earth orbit (GEO), with less focus on polar Earth orbit (PEO)], as well as spacecraft in other energetic plasma environments, such as those at Jupiter and Saturn, and interplanetary solar wind charging environments. Designs for spacecraft with orbits in these regions should be evaluated for the threat of external (surface) and/or internal charging, as noted. NASA RP-1354, Spacecraft Environments Interactions: Protecting Against the Effects of Spacecraft Charging (2), describes environmental interaction mitigation design techniques at an introductory level.
Specifically, this book does not address LEO spacecraft charging at orbital inclinations such that the auroral zones are seldom encountered. That region is the purview of NASA-STD-4005 (3) and NASA-HDBK-4006 (4). The book is intended to be complementary to those standards and applies to other regions. In particular, mitigation techniques for low-inclination LEO orbits may differ from those that apply to regions covered by this book. Spacecraft in orbits, such as GEO transfer orbits that spend time in both regimes, should use mitigation techniques that apply to both regimes. It also does not include such topics as the following:
Figures 1.1 and 1.2 illustrate the approximate regions of concern for charging as defined in this book. Figure 1.1 is to be interpreted as the worst-case surface charging that may occur in the near-Earth environment. The north/south latitudinal asymmetry assumes that the magnetic North Pole is tilted as much as possible for this view. Potentials are calculated for an aluminum sphere in shadow. Note that at altitudes above 400 km, spacecraft charging can exceed 400 to 500 V, which has the possibility of generating discharges. Indeed, the Defense Meteorological Satellite Program (DMSP) and other satellites have reported significant charging in the auroral zones many times (as high as -4000 V), and one satellite [Advanced Earth Observation Satellite II (ADEOS-II)] at 800 km experienced total failure due to spacecraft charging (5–7).
Figure 1.2, which illustrates Earth's internal charging threat regions, is estimated assuming averages over several orbits since the internal charging threat usually has a longer time scale and reflects the approximate internal charging threat for satellites with the indicated orbital parameters. It is intended to illustrate the approximate regions of concern for internal electrostatic discharge (IESD).
In this book, the distinction between surface charging and internal charging is that internal charging is caused by energetic particles that can penetrate and deposit charge very close to a victim site. Surface charging occurs on areas that can be seen and touched on the outside of a spacecraft. Surface discharges occur on or near the outer surface of a spacecraft, and discharges must be coupled to an interior affected site rather than directly to the victim. Energy from surface arcs is attenuated by the coupling factors necessary to get to victims (most often inside the spacecraft) and therefore is less of a threat to electronics. External wiring and antenna feeds, of course, are susceptible to this threat. Internal charging, by contrast, may cause a discharge directly to a victim pin or wire with very little attenuation if caused by electron deposition in circuit boards, wire insulation, or connector potting.
Geosynchronous orbit (a circular orbit in the equatorial plane of Earth at about 35,786 km altitude) is perhaps the most common example of a region where spacecraft are affected by spacecraft charging, but the same problem can occur at lower Earth altitudes, in Earth polar orbits, at Jupiter, and at other places where spacecraft can fly. Internal charging is sometimes called deep dielectric charging or buried charging. Use of the word dielectric can be misleading, since ungrounded internal conductors can also present an internal electrostatic discharging threat to spacecraft. This book details the methods necessary to mitigate both in-flight surface and internal charging concerns as the physics and design solutions for both are often similar.
1. R. D. Leach and M. B. Alexander, Eds., Failures and Anomalies Attributed to Spacecraft Charging, NASA Reference Publication 1375, National Aeronautics and Space Administration, August 1995. This document has a very good list of specific space incidents that have been attributed to electrostatic discharges in space. It does not discriminate between surface charging or internal charging, but that is usually difficult to determine or does not appear in public literature.
2. J. L. Herr and M. B. McCollum, Spacecraft Environments Interactions: Protecting Against the Effects of Spacecraft Charging, NASA-RP-1354, National Aeronautics and Space Administration, 1994.
3. D. C. Ferguson, Low Earth Orbit Spacecraft Charging Design Standard, NASA-STD-4005, 16 pages, National Aeronautics and Space Administration, June 3, 2007.
4. D. C. Ferguson, Low Earth Orbit Spacecraft Charging Design Handbook, NASA-HDBK-4006, 63 pages, National Aeronautics and Space Administration, June 3, 2007.
5. D. L. Cooke, “Simulation of an Auroral Charging Anomaly on the DMSP Satellite,” 36th Aerospace Sciences Meeting and Exhibit, Reno, Nevada, AIAA-98-0385, January 12–15, 1998.
6. S. Kawakita, H. Kusawake, M. Takahashi, H. Maejima, J. Kim, S. Hosoda, M. Cho, K. Toyoda, and Y. Nozaki, “Sustained Arc Between Primary Power Cables of a Satellite,” 2nd International Energy Conversion Engineering Conference, Providence, Rhode Island, August 16–19, 2004. Contains description of ADEOS-II satellite failure analysis. See also Maejima et al. (7).
7. H. Maejima, S. Kawakita, H. Kusawake, M. Takahashi, T. Goka, T. Kurosaki, M. Nakamura, K. Toyoda, and M. Cho, “Investigation of Power System Failure of a LEO Satellite,” 2nd International Energy Conversion Engineering Conference, Providence, Rhode Island, August 16–19, 2004. Contains description of ADEOS-II satellite failure analysis.
8. R. W. Evans, H. B. Garrett, S. Gabriel, and A. C. Whittlesey, “A Preliminary Spacecraft Charging Map for the Near Earth Environment,” Spacecraft Charging Technology Conference, Naval Postgraduate School, Monterey, California, November 1989. This original reference paper was omitted from the conference proceedings. See Whittlesey et al. (9) for an alternative reference with the “wishbone” chart.
9. A. Whittlesey, H. B. Garrett, and P. A. Robinson, Jr., “The Satellite Space Charging Phenomenon, and Design and Test Considerations,” IEEE International EMC Symposium. Anaheim, California, 1992.
Chapter 2
Introduction to the Physics of Charging and Discharging
The fundamental physical concepts that account for space charging are described in this chapter. The appendices expand this description by means of equations and examples.
Spacecraft charging occurs when charged particles from the surrounding plasma and energetic particle environment stop on the spacecraft: either on the surface, on interior parts, in dielectrics, or in conductors. Other items affecting charging include biased solar arrays or plasma emitters. Charging can also occur when photoemission occurs; that is, solar photons cause surfaces to emit photoelectrons. Events after that determine whether or not the charging causes problems.
A plasma is a partially ionized gas in which some of the atoms and molecules that make up the gas have some or all of their electrons stripped off, leaving a mixture of ions and electrons that can develop a sheath that can extend over several Debye lengths. Except for LEO, where ionized oxygen (O+) is the most abundant species, the simplest ion, a proton (corresponding to ionized hydrogen, H+), is generally the most abundant ion in the environments considered here. The energy of the plasma, its electrons and ions, is often described in units of electron volts (eV). This is the kinetic energy that is given to the electron or ion if it is accelerated by an electric potential of that many volts. Whereas temperature (T) is generally used to describe the disordered microscopic motion of a group of particles, plasma physicists also use it as another unit of measure to describe the kinetic energy of the plasma. For electrons, numerically T(K) equals T(eV) × 11,604; that is, 4300 eV is equivalent to 50 million degrees kelvin (K).
The kinetic energy of a particle is given by the equation
2.1
where
E = energy
m = mass of the particle
v = velocity of the particle
Because of the difference in mass ( ∼ 1 : 1836 for electrons to protons), electrons in a plasma in thermal equilibrium generally have a velocity about 43 times that of protons. This translates into a net instantaneous flux or current of electrons onto a spacecraft that is much higher than that of the ions [typically, nanoamperes per square centimeter (nA/cm2) for electrons versus picoamperes per square centimeter (pA/cm2) for protons at geosynchronous orbit]. This difference in flux is one reason for the charging effects observed (a surplus of negative charges on affected regions). For electrons, numerically the velocity (ve) equals km/s and for protons the velocity (vp) equals km/s, when E is in eV.
Although a plasma may be described by its average energy, there is actually a distribution of energies. The rate of charging in the interior of the spacecraft is a function of the flux versus energy, or spectrum, of the plasma at energies well in excess of the mean plasma energies [for GEO, the mean plasma energy may reach a few tens of kiloelectron volts (keV)]. Surface charging is usually correlated with electrons in the approximate 0 to 50 keV energy range, while significant internal charging is associated with the high-energy electrons [100 keV to 3 megaelectron volts (MeV)].
A simple plasma and its interactions with a surface are illustrated in Figs. 2.1 and 2.2. The electrons (e−) and ions (represented by H+ in Fig. 2.2) are moving in random directions (omnidirectional) and with different speeds (a spectrum of energies). Figure 2.2 illustrates surface charging. (Exterior surfaces are shown; the interior is similar.) To estimate surface charging, both the electron and ion spectra should be known from about 1 eV to 100 keV. Although fluxes might be directed, omnidirectional fluxes are assumed in this document because spacecraft orientation relative to the plasma is often not well defined.
Electrons and ions will penetrate matter. The depth of penetration of a given species (electron, proton, or other ion) depends on its energy, its atomic mass, and the composition of the target material. Figure 2.3 shows the mean penetration range versus the energy of electrons and protons into aluminum and represents the approximate penetration depth into a slab of aluminum. To first order, only particles with an energy corresponding to a range greater than the spacecraft shield thickness can penetrate the spacecraft interior. If the material is not aluminum, an equivalent penetration depth is roughly the same number of grams per square centimeter of the material's thickness.
In this document we use the terms surface charging and internal charging. In the literature the terms buried dielectric charge or deep dielectric charge for internal charging are also used, but these terms are misleading because they give the impression that only dielectrics can accumulate charge. Although dielectrics can accumulate charge and discharge to cause damage, ungrounded conductors can also accumulate charge and must also be considered an internal charging threat. In fact, ungrounded conductors can discharge with a higher peak current and a higher rate of change of current than a dielectric and can be a greater threat.
Based on typical spacecraft construction, there is usually an interior section, referred to in this document as internal. It is assumed that this interior section has shielding of at least 3 mils of aluminum equivalent, corresponding to electron energies greater than 0.1 MeV. Surface charging would be the outer layers of the spacecraft, corresponding to 2 mils of aluminum or 0 to 50 keV electrons. Obviously, the surface/internal charging cutoff depends on spacecraft construction. Protons are often not considered for spacecraft charging because the greater impinging flux of electrons at the same energy and (for internal charging) the lesser penetration of protons reduces the internal flux to a negligible amount. Higher atomic mass particles are even less of a threat because of their much lower fluxes.
Because electrons may stop at a depth that is less than their maximum penetration depth and because the electron spectrum is continuous, the penetration depth/charging region will be continuous, ranging from the charges deposited on the exterior surface to those deposited deep in the interior. Internal charging as used here often is equivalent to “inside the Faraday cage.” For a spacecraft that is built with a Faraday cage thickness of 30 or more mils of aluminum equivalent, this would mean that internal effects deal with the portion of the electron spectrum above 500 keV and the proton spectrum above 10 MeV. At GEO orbits, the practical range of energy for internal charging is 100 keV to about 3 MeV, bounded on the lower end by the fact that most spacecraft have at least 3 mils of shielding and on the upper end by the fact that, as will be shown later, common GEO environments above 3 MeV do not have enough plasma flux to cause internal charging problems.
Figure 2.4 illustrates the concept that energetic electrons will penetrate interior portions of a spacecraft. Having penetrated, the electrons may be stopped in dielectrics or on ungrounded conductors. If too many electrons accumulate, the resulting high electric fields inside the spacecraft may cause an electrostatic discharge (ESD) to a nearby victim circuit. Note that the internal charging resembles surface charging, with the exception that circuits are rarely exposed victims on the exterior surface of a spacecraft, and thus (with the condition that charging rates are slower) internal charging results in a greater direct threat to circuits.
The term ESD in this document is general or may refer to surface discharges. The term internal ESD (IESD) refers to ESDs on the interior regions of a spacecraft as defined above.
The first step in analyzing a design for the internal charging threat is to determine the charge deposition inside the spacecraft. It is important to know the amount of charge deposited in or on a given material, as well as the deposition rate, as these determine the distribution of the charge and hence the local electric fields. An electrical breakdown (discharge) will occur when the local electric field exceeds the dielectric strength of the material or between dissimilar surfaces with a critical potential difference. The actual breakdown can be triggered by a variety of mechanisms including the plasma cloud associated with a micrometeoroid or space debris impact. The amplitude and duration of the resulting pulse are dependent on the charge deposited. These values in turn determine how much damage may be done to spacecraft circuitry.
Charge deposition is a function not only of the spacecraft configuration but also of the external electron spectrum. Given an electron spectrum and an estimate of the exterior shielding, the penetration depth versus the energy chart (Fig. 2.3) permits an estimate of electron deposition as a function of depth for any given equivalent thickness of aluminum, from which the likelihood of a discharge can be predicted. Because of such complexities as hardware geometries, however, it is normally better to run an electron penetration or radiation shielding code to more accurately determine the charge deposited at a given material element within a spacecraft. Appendixes B and C list some environment and penetration codes.
Material conductivity plays an important role in determining the likelihood of a breakdown. The actual threat posed by internal charging depends on accumulating charge until the resulting electric field stress causes an ESD. Charge accumulation depends on retaining the charge after deposition. Since internal charging fluxes at GEO are on the order of 1 pA/cm2 (1 pA = 10−12 A), resistivities on the order of 1012 Ω · cm will conduct charge away, if grounded, so that high local electric field stress (105 to 106 V/cm) conditions cannot occur and initiate an arc. Unfortunately, modern spacecraft dielectric materials such as Teflon® and Kapton®, flame retardant 4 (FR4) circuit boards, and conformal coatings often have high enough resistivities to cause problems (Section 6.1). If the internal charge-deposition rate exceeds the leakage rate, these excellent dielectrics can accumulate charge to the point that discharges to nearby conductors are possible. If that conductor leads to or is close to a sensitive victim, there could be disruption or damage to the victim circuitry.
Metals, although conductive, may be a problem if they are electrically isolated by more than 1012 Ω. Some examples of metals that may be isolated (undesirable) are radiation spot shields, structures that are deliberately insulated, capacitor cans, integrated-circuit (IC) and hybrid cans, transformer cores, relay coil cans, and wires that may be isolated by design or by switches. Each of these isolated items could be an internal charging threat and should be scrutinized for its contribution to the internal charging hazards.
The breakdown voltage is that voltage at which the dielectric field strength of a particular sample (or air gap) cannot sustain the voltage stress and a breakdown (arc) is likely to occur. The breakdown voltage depends on the basic dielectric strength of the material [volts per mil (V/mil) is one measure of the dielectric strength] and on the thickness of the material. Even though the dielectric strength is implicitly linear, the thicker materials usually are reported to have less strength per unit thickness. Manufacturing blemishes or handling damage can all contribute to the variations in breakdown strength that will be observed in practice. As a rule of thumb, if the exact breakdown strength is not known, most common good-quality spacecraft dielectrics may break down when their internal electric fields exceed 2 × 105 V/cm (2 × 107 V/m; 508 V/mil). As a practical matter, because of sharp corners, interfaces, and vias that are inevitably present in printed circuit (PC) boards, the breakdown voltage may be less.
The dielectric constant of a material, or its permittivity, is a measure of the electric field inside the material compared to the electric field in a vacuum. It is commonly used in the description of dielectric materials. The dielectric constant of a material ( ε ) is generally factored into the product of the permittivity of free space ( ε 0 = 8.85 × 10−12 F/m) and the relative permittivity ( ε r, a dimensionless quantity) of the material in question ( ε = ε 0 × ε r). Relative dielectric constants of insulating materials used in spacecraft construction generally range from 2.1 to as much as 7: assuming a relative dielectric constant of 2.7 (between Teflon® and Kapton®) is an adequate approximation if the exact dielectric constant is not known. Examples of the use of the dielectric constant for calculating time constants are given in Section E.7.
The density of a material is important in determining its shielding properties. The penetration depth of an electron of given energy, and therefore its ability to contribute to internal charging, depends on the thickness and density of the material through which it passes. Since aluminum is a typical material for spacecraft outer surfaces, the penetration depth is commonly based on the aluminum equivalent. To the first order, the penetration depth in materials depends on the shielding mass. That is, if a material is one-half the density of aluminum, it takes twice the thickness to achieve the same shielding as that of aluminum.
For IESD, the electron flux for a given duration at a location is a critical quantity. Figure 2.5 compares spacecraft disruptions as functions of environmental flux at the victim location. Experience and observations from the Combined Release and Radiation Effects Satellite (CRRES) and other satellites have shown that if the normally incident internal flux is less than 0.1 pA/cm2, there have been few, if any, internal charging problems [2 × 1010 electrons/cm2 in 10 h appears to be the threshold]. Bodeau (1, 2) and others report problems with sensitive circuits at even lower levels on some newer spacecraft. For geosynchronous orbits, the flux above 3 MeV is usually less than 0.1 pA/cm2, and a generally suitable level of protection can be provided by 110 mils of aluminum equivalent (Fig. 2.3). Modern spacecraft are being built with thinner walls or only thermal blankets (less mass), so the simple solution to the internal charging problem (adding shielding everywhere) cannot be implemented. However, adding spot shielding mass (grounded) near sensitive regions can help in many cases.
Figure 2.5 also allows a direct comparison between common units as used in the literature (i.e., 106 e/cm2 · s is about 0.2 pA/cm2). Additional information about CRRES is provided in Section B.1.2.5.
The approximation of 0.1 pA/cm2 noted as a nominal threshold for internal charging difficulties is experientially based, not physics based, and thus has limits. Some considerations include that this is based on CRRES data (although verified by other researchers) for “typically used materials” and probably at or near room temperature. If highly resistive materials are used in cold situations and near electronics, further testing or analysis should be done.
To assess the magnitude of the IESD concern for a given orbit, it is necessary to know the electron charging environment along that orbit. (As noted earlier, the protons generally do not have enough penetrating flux to cause a significant internal charge.) The electron orbital environments of primary interest (in terms of number of affected satellites) are GEO, medium Earth orbits (MEOs), and polar Earth orbits (PEOs). Other orbital regimes that are also known to be of interest are Molniya orbits and orbits at Jupiter and Saturn (Sections B.3 and B.4).
The 11-year variation between the most severe electron environments and the least severe can vary over a 100 : 1 range and shows a correlation with the solar cycle (Section B.2.2.1, Figs. B-3 and B-4). A project manager might consider “tuning” the protection to the anticipated service period, but even in quiet years, the worst flux will sometimes be as high as the worst flux of noisy years. The environment presented in this document represents a worst-case level for GEO for any phase of the solar cycle.
Figure 2.6 shows a worst-case GEO internal charging spectrum generated by selecting dates when the Geosynchronous Operational Environmental Satellite (GOES) EFigure 2.6°