NASATechnical, Paper2361

1984

Na tlonal Aeronauticsand Space AdmInIstratlon

Scientific and TechnicalInformation Branch

Design Guidelines for ‘

*

Assessing and ControllingSpacecraft Charging Effects

Carolyn K. PurvisLewis Research CenterCleveland, Ohio

Henry B. Garrettand A. C. WhittleseyCalijknia Institute of TechnologyJet Propulsion LaboratoryPasadena, Calijornia

N. John StevensHughes Aircraft CompanyEl Segundo, Calfornia

ForewordExperience has indicated a need for uniform criteria, or guidelines, to be used in

all phases of spacecraft design. Accordingly, guidelines have been developed for thecontrol of absolute and differential charging of spacecraft surfaces by the lowerenergy (less than approximately 50 keV) space charged-particle environment.Interior charging due to higher energy particles was not considered.

This document is to be regarded as a guide to good design practices for assessingand controlling charging effects. It is not a NASA or Air Force mandatory require-ment unless specifically included in project specifications. It is expected, however,that this document, revised as experience may indicate, will provide uniform designpractices for all space vehicles.

The guidelines have been compiled from published information by the NASALewis Research Center and the California Institute of Technology, Jet PropulsionLaboratory. Comments concerning the technical content of this document should beaddressed to C. K. Purvis at the NASA Lewis Research Center.

Some of the information contained in this document was assembled undercontract NAS3-21048 by Robert E. Kamen and Alan B. Holman, SAI l

Incorporated. Significant contributions to that effort were made by EdwardO’Donnell, Michael Grajek, Rita Simas, and Donald McPherson, SAIIncorporated. Special appreciation is extended to the following companies andagencies for the information they supplied for this document: Air Force MaterialsLaboratory, Beers Associates, Communication Spacecraft Corporation, FordAerospace, General Electric, Hughes Aircraft, IRT Corporation, NASA GoddardSpace Flight Center, Naval Research Laboratory, Mission Research Corporation,

. Rockwell-International, and the Air Force Space Division. K. Duff and D. Hoshinoprepared the manuscript and G. Plamp provided data on material properties.

Contents. Page

1 .O Introduction1.1 Definition of Spacecraft Charging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2 Spacecraft Charging Concerns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.3 Initial Environmental Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.3.1 Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.3.2 Spacecraft role . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.3.3 Spacecraft configuration. . . . . . . . . . . . . . . . . ..*................................ 21.3.4 Effect of charging on systems . . . . . . . . . . . . . . . . . . . . . . . . ..-................... 2

1.4 Design Guidelines Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.0 Spacecraft Modeling Techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2.1 Substorm Environment Specifications . . . . . . . ..m................................. 32.2 Spacecraft Surface Charging Models . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2.2.1 Simple approximations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..~............... 32.2.2 NASA Charging Analyzer Program (NASCAP) . . . . . . . . . . . . . . . . . . . . . . 5

2.3 Discharge Characteristics . . . . . . . . . . . . . ..**........*................................. 52.3.1 Dielectric surface breakdowns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52.3.2 Buried charge breakdowns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72.3.3 Spacecraft-to-space breakdowns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.4 Coupling Models . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82.4.1 Lumped-element modeling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82.4.2 Specification and Electromagnetic Compatibility

Program (SEMCAP) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 .O Spacecraft Design Guidelines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.1 General Guidelines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113.1.1 Grounding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113.1.2 Exterior surface materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113.1.3 Shielding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153.1.4 Filtering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..*................................... 153.1.5 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . ..*...............*........................ 15

3.2 Subsystem Guidelines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153 -2.1 Electronics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153.2.2 Power systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163.2.3 Mechanical and structural . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163.2.4 Thermal control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163.2.5 Communications systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173.2.6 Attitude control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173.2.7 Payloads . . . . . . . . . . . . . . . ..*.......*.............................................. 17

4.0 Spacecraft Test Techniques ............................................................. 184.1 Test Philosophy ....................................................................... 184.2 Simulation of Parameters ........................................................... 184.3 General Test ‘Methods ............................................................... 19

4.3.1 ESD-generating equipment ................................................. 194.3.2 Methods of ESD application ................................................ 21

4.4 Unit Testing............................................................................ 224.4.1 General ......................................................................... 224.4.2 Unit test configuration 23.......................................................4.4.3 Unit test operating modes ................................................... 23

4.5 Spacecraft Testing .................................................................... 234.5.1 General ......................................................................... 234.5.2 Spacecraft test configuration .............................................. 23

5.0 Control and Monitoring Techniques .................................................. 255.1 Active Spacecraft Charge Control ................................................ 255.2 Environmental and Event Monitors .............................................. 25

Appendix A. Description of Geosynchronous Plasma Environments .............. 26Appendix B. Technical Description of NASCAP ....................................... 31Appendix C. Voyager SEMCAP Analysis ................................................ 34Bibliography..................................................................................... 36

1.0 Introduction spacecraft surfaces or between spacecraft surfaces andsDacecraft ground. When a breakdown threshold is ex-

These guidelines are intended to provide a ready ceeded, an electrostatic discharge can occur. Thereference for spacecraft systems designers and others transient generated by this discharge can couple into theneeding an overall view of the techniques required to spacecraft electronics and cause upsets ranging fromlimit the detrimental effects of spacecraft charging. The logic switching to complete system failure. Dischargesprimary goals of this document are to summarize the can also cause long-term degradation of exterior surfaceavailable information *on controlling charging effects and coatings and enhance contamination of surfaces. Vehicleto provide guidelines for incorporating immunity to torquing or wobble can also be produced when multipleelectromagnetic transients into spacecraft and spacecraft discharges occur. The ultimate results are disruptions insubsystems. spacecraft operation.

1.1 Definition of Spacecraft ChargingSpacecraft charging is defined as those phenomena

associated with the buildup of charge on exposed externalsurfaces of geosynchronous spacecraft. This surfacecharging results from spacecraft encounter with ageomagnetic substorm environment-a plasma withparticle energies from 1 to 50 keV.

Two types of spacecraft charging will be considered.The first, called absolute charging, occurs when the entirespacecraft potential relative to the ambient space plasmais changed uniformly by the encounter with the chargingenvironment. The second type, called differential charg-ing, occurs when parts of the spacecraft are charged todifferent negative potentials relative to each other. In thistype of charging, strong local electric fields may exist.

1 .Z Spacecraft Charging ConcernsThe designer must recognize the importance of mission

role and spacecraft configuration in evaluating absoluteand differential charging effects. The buildup of largepotentials on spacecraft relative to the ambient plasma isnot, of itself, a serious electrostatic discharge (ESD)design concern. However, such charging enhances sur-face contamination, which degrades thermal properties.It also compromises scientific missions ‘seeking tomeasure properties of the space environment. Spacecraftsystems referenced to structure ground are not affectedby a uniformly charged spacecraft. However, spacecraftsurfaces are not uniform in their material properties,surfaces will be either shaded or sunlit, and the ambientfluxes may be anisotropic. These and other chargingeffects can produce potential differences between

1.3 Initial Environmental Considerations1.3.1 Environment

be overtaken by the hot plasma. The problem for the used and since sunlight can illuminate only one side at aspacecraft designer is that each of these environments time, there will always be some differential charging asrepresents a unique set of plasma conditions as viewed by well as absolute charging. The effect of this surfacethe spacecraft and results in a markedly different charging on the per t’ortlxlncc of spn~t~a ft must becharging history. evaluated in t ems of r\x~Ifun&~~lS, ~pscts, irnd t’iriltws.

For absolute charging the spacecraft potential changesas a whole-the dielectric surface voltages are “locked”to the ground reference voltage. This type of chargingoccurs very rapidly (in fractions of a second), typicallyduring eclipse. Differential charging usually occursslowly (in minutes) and results in one part or surfacebeing charged to a potential different from those of otherparts of the spacecraft. This differential charging canalso change the absolute charging level of the spacecraft.This is the usual mechanism for daylight charging, whichconsequently occurs slowly.

As stated, surface charpinp r’~ul~i ~!isrupt cw it-w -mental measurements on s4zientitk spacccrati . For thisapplication and others where control of electrostaticfields is required, material selection to minimizedifferential charging is mandatory. For operationalspacecraft, surface charging can also cause problems.The hallmark of the spacecraft charging phenomenon isthe occurrence of electronic switching anomalies. Theseanomalies are believed to result from transients causedby differential-charging-induced discharges. Theseanomalous events even seem to occur in systems that aresupposedly immune to noise. The discharge-inducedtransients, under very severe environmental conditions,can cause system failures.

1.3.2 Spacecraft roleA critical factor influencing the extent to which charg-

ing interactions must be controlled is the mission of thespacecraft. In all spacecraft, differential charging isundesirable. For scientific spacecraft, absolute chargingusually is not desired. For such spacecraft, conductive-coated dielectrics can be used to minimize differentialsurface charging, and active charge control devices can beincorporated to hold the spacecraft potential close to thespace plasma potentiai. For operational spacecraft theeffort should be directed toward controlling thosecharging effects that are detrimental to the particular

Surface charging also enhances contamination. Thecontaminants are attracted back to charged surfaces and ’deposit on them. This changes surface characteristics.Altered surface optical properties result in higher tern- .peratures. Changes in secondary and photoelectron yieldsresult in altered charging characteristics. Deposition ofdielectric contaminants can also reduce surfaceconductivity. If there are severe discharges on thesurfaces, the materials can be damaged and can changethe thermal control performance.

mission.More definitive data on environmentally induced

1.4 Design Guidelines Format

effects in geosynchronous satellites are needed, This database could be obtained if all such spacecraft carriedenvironment and event monitors (section 5.2).

1.3.3 Spacecraft configurationAlso of major concern in determining the importance

of spacecraft charging is the effect of spacecraft config-uration on charging behavior. A spin-stabilizedspacecraft usually has a low spacecraft ground potential(a few hundred volts negative). On some shaded dielectricsurfaces during sunlit charging events, differentialvoltages of several thousand volts can occur.

A three-axis-stabilized spacecraft can have a ratherlarge negative structure potential (a few thousand volts)in sunlit charging events. The dominant areas controllingcharging in this case are the backs of the solar arraywings. Differential charging will likely not be as large asin the spinner case (Purvis, 1980).

1.3.4 Effect of charging on systemsThe geosynchronous substorm environment will charge

spacecraft exterior surfaces. Since different materials are

This document has been prepared as a guide tospacecraft system designers. These guidelines should beused early in the design process so that the control ofspacecraft charging can be easily and economicallyachieved. It should be stressed that, if such control is tobe successfully incorporated, care must be exercisedthroughout the program to ensure compliance with theguidelines. Each spacecraft is different, and thesegeneralized guidelines must be adapted and modified tofit the particular application. .

The document is divided into five parts. The firstsection has introduced spacecraft charging concepts ofimportance to the designer. The following section detailsthe modeling techniques to be used to assess whether thedesign is adequate for environmental immunity. Thethird section presents specific guidelines for protectingsystems and subsystems. This is followed by a sectiondescribing test procedures for demonstrating systemimmunity. The fifth section discusses active chargecontrol and monitoring techniques. Appendixes presentillustrative examples and the bibliography lists otherdocuments for those desiring further information onspecific topics.

2.0 Spacecraft Modeling Techniques

. Modeling is an essential activity in spacecraft designand in evaluating spacecraft charging effects. There arefour regimes of interest in modeling these effects. First,the ambient environment and its fluctuations must bespecified. Second, the interaction process-the buildupof charge and electric fields near the vehicle-must bemodeled. Third, given the existence of charged surfacesand potential gradients, the likelihood, signal charac-teristics, and frequency of electrostatic discharge must bemodeled. Finally, the coupling of the electrostaticdischarge pulse to individual circuit elements must bemodeled in order to identify the spacecraft elements mostlikely to be affected. Recommended modeling proceduresare presented in this section along with overviews of thephysical processes involved.

For brevity in the discussion that follows, some of themore detailed material has been placed in the appendixes.

2.1 Substorm Environment Specifications

Maxwellian temperatures are presented-and these onlyfor the electrons and protons. Useful answers can beobtained with this simple representation. For a worst-case static charging analysis the “single Maxwellian”environmental characterization given in table I iorecommended.

The values given in table I are a 90th percentile single-Maxwellian representation of the environment (appendixA). Section 2.3 describes the spacecraft chargingequations and methods in which these values will be usedto predict spacecraft charging effects. If the worst-caseanalysis shows that spacecraft surface differentialpotentials are less than 500 V, there should be noelectrostatic discharging problem. If the worst-caseanalysis shows a possible problem, use of more realisticplasma parameters should be considered.

A more comprehensive discussion of plasmaparameters is given in appendix A. Some original data arepresented for the ATS-5, ATS-6, and SCATHAsatellites, with average values, standard deviations, andworst-case values. Additionally, percentages of yearlyoccurrences are given, and finally, a typical time historyof a model substorm is shown. All of these differentdescriptions of plasma parameters can be used to helpanalyze special or extreme spacecraft charging situations.

2.2 Spacecraft Surface Charging ModelsAnalytical modeling techniques should be u&d to

predict surface charging effects. In this section,approaches to predicting spacecraft surface voltagesresulting from encounters with the substorm environmentare discussed. The predictions identify possible dischargelocations and are used to establish the spacecraft andcomponent level test requirements.

2.2.1 Simple approximationsThe simple approximations discussed in this section are

of a worst-case nature. If. this analysis indicatesdifferential potentials of less than 500 V, there should beno spacecraft discharge problems. If predicted potentialson materials exceed 500 V, the NASA Charging AnalyzerProgram (NASCAP) code (section 2.2.2) must be used.

Although the physics behind the spacecraft chargingprocess is quite complex, the formulation at geosyn-chronous orbit can be expressed in very simple terms if a

TABLE I. - WORST-CASE GEOSYNCHRONOUS PLASMAENVIRONMENT

4Electron number density, NE, cm-3 . 1.12Electron temperature, TE, es

............1.2~10~

Ion number density, NI, cm- . . . . . . 2.36x10-1Ion temperature, TI, eV . . . . . . . . . 2.95x104

Maxwell-boltzmann distribution is assumed. Thefundamental physical process for all spacecraft chargingis that of current balance-at equilibrium, all currentsmust sum to zero. The potential at which equilibrium isachieved is the potential difference between thespacecraft and the space plasma ground. The basicequation expressing this current balance for a givensurface in an equilibrium situation is, in terms of thecurrent:

V spacecraft potentialIE incident electron current on spacecraft surfaceII incident ion current on spacecraft surfacekE secondary electron current due to ZE

.IS1 secondary electron current due to ZlZBSE backscattered electrons due to ZEZPH photoelectron currentIB active current sources such as charged particle

beams or ion thrustersJT total current to spacecraft (at equilibrium, ZT=O)

For a spherical body and a Maxwell-Boltzmanndistribution, the first-order current densities (the currentdivided by the area over which the current is collected)can be shown (Garrett, 1981) to be given by

Electrons

V c 0 repelled

V>O attracted

Ions

V > 0 repelled\

IWO attracted

where

JEo= (9) ( $)1’2

and

JIo= ($!) (z)1’2

,

4

(2)

(3)

(4)

(9

(6)

where NE and s NI are densities of electrons and ions,respectively; rn~ and rnr are masses of electrons and ions,respectively; and q is the magnitude of the electroniccharge.

Given these expressions and parameterizing the

.

secondary and backscatter emissions, equation (1) can bereduced to an analytic expression in terms of the potentialat a point. This model, called an analytic probe model,can be stated as follows:

-APHJPH&X~)=ZT=O Ye0 (7)

where

AE electron collection areaIJEO ambient electron current density4 ion collection areaJlO ambient ion current densityAPH photoelectron emission areaJPHO . saturation photoelectron fluxBSE,SE,SI parameterization functions for secondary

emission due to backscatter, electrons, andions

f(xm) attenuated solar flux as a function ofaltitude Xm of center of Sun above surfaceof Earth as seen by spacecraft, percent

This equation is appropriate for a small (< 10 m),uniformly conducting spacecraft at geosynchronous orbitin the absence of magnetic field effects. To solve theequation, V is varied until ZT=O. Typical values of SI,SE, and BSE are 3, 0.4, and 0.2, respectively, foraluminum. For geosynchronous orbit, J&J1 is about 30during a geomagnetic storm. When the spacecraft is ineclipse, these values give

Vz-TE (8)

where TE is in electron volts. That is, to first order ineclipse; the spacecraft potential is approximatelynumerically equal to the plasma temperature expressed inelectron volts. Note, however, that TE must exceed somecritical value (Olsen, 1983; Garrett et al., 1979), usuallyof the order of 1000 eV, before charging will occurbecause secondary electron production can ‘exceedambient current for low enough TE.

2.2.2 NASA Charging Analyzer Program (NASCAP)

.

The NASA Charging Analyzer Program computercode (Katz et al., 1977, 1979; Schnuelle et al., 1979;Roche and Purvis, 1979; Rubin et al., 1980) has beenspecifically developed as an engineering tool to determinethe environmental effect on spacecraft surfaces andsystems. It can analyze the surface charging of a three-dimensional, complex body as a fun&on of time forgiven space environmental conditions and specifiedsurface potentials. Material properties of surfaces areincluded in the computations. Surface potentials, low-energy sheath properties, potential distributions in space,and particle trajectories are computed. By locating severesurface voltage gradients in a particular design, it ispossible to show where discharges could occur. The effectof changes in the surface materials or coatings in thoseareas on minimizing voltage gradients can then beevaluated. The environment of table I should be used inthese analyses. NASCAP is described in detail inappendix B.

2.3 Discharge CharacteristicsCharged spacecraft surfaces can discharge, and the

resulting transients can couple into electrical systems. Aspacecraft in space must be considered to be a capacitorrelative to the space plasma potential. The spacecraft, inturn, is divided into numerous other capacitors by thedielectric surfaces used for thermal control and for powergeneration. This system of capacitors can be charged atdifferent rates depending on incident fluxes, time con-stants, and spacecraft configuration effects. Because ofthis complex charging rate pattern, sophisticatedcomputer programs are required to predict behavior.

The system of capacitors floats electrically with respectto the space plasma potential. This can give rise tounstable conditions in which charge can be lost from thespacecraft to space. Whether anyone will ever be able toestablish exact conditions required for such breakdownsis questionable. What is known is that breakdowns dooccur, and it is hoped that conditions that lead to break-downs can be bounded.

Breakdowns, or discharges, probably occur becausea differential charge builds up in spacecraft dielectricsurfaces or between various surfaces on the spacecraft.Whenever this charge buildup generates an electric fieldthat exceeds a breakdown threshold, charge will bereleased from the spacecraft to space. This charge releasewill continue until the differential driving force no longerexists. Hence, the amount of charge released will belimited to the total charge stored in or on the dielectric atthe discharge site. The charge loss or current to spacecauses the dielectric surface voltage (at least locally) torelax toward zero. Since the dielectric is capacitivelycoupled to the structure, the charge loss will also causethe structure potential to become less negative. In fact,

2.3.1 Dielectric surface breakdowns

If either of thedischarges can occur:

following criteria are exceeded,

(I) Dielectric surface voltages are greater than 500 Vpositive relative to an adjacent exposed conductor.

(2) The interface between a dielectric and an exposedconductor has an electric field greater than 1 x 105 V/cm.Note that edges, points, gaps, seams, and imperfectionsin surface materials can increase electric fields and hencepromote the probability of discharges. These items arenot usually modeled and must be found by close inspec-tion of the exterior surface specifications.

The first criterion can be exceeded by solar arrays inwhich the high secondary yield of the coverslide canresult in surface voltages that are positive with respect tothe metalized interconnects. This criterion can also apply

to metalized dielectrics in which the metalized film, eitherby accident or design, is isolated from structure groundby a resistance value greater than 10 MQ (essentially onlycapacitively coupled). In the latter case, the dielectric canbe charged to large negative voltages (when shaded), andthe metal film will thus become more negative than thesurrounding surfaces and act as a cathode or electronemitter.. : The second criterion applies to those areas of aspacecraft where a strong negative voltage gradient could

* exist. This usually would be associated with the edge of adielectric next to another surface or with cracks in thedielectric exposing a conductor underneath. The chargestored on or in the dielectric is relatively unstable andcould be lost.: In both of these conditions, stored charge is initiallyejected to space in the discharge process. This lossproduces a transient that can couple into the spacecraftIstructure and possibly into the electronic systems.Current returns from space to the exposed conductiveareas of the spacecraft. Transient currents flow in thestructure depending on the electrical characteristics. It isassumed that the discharge process will continue until thevoltage gradient or electric field that began the processdisappears. The currents flowing in the structure will

: damp out according to its resistance.The computation of charge lost in any discharge is

highly speculative at this time. Basically, charge loss canbe considered to result from the depletion of twocapacitors: that stored in the spacecraft, which is chargedto a specified voltage relative to space, and that stored ina limited region of the dielectric at the discharge site. Thecomputation of the charge loss is a question of judgmenton the part of the analyst and must depend on thepredicted voltages on the spacecraft at the time thatdischarges are expected to occur. As a guide the followingcharge loss categories might be useful:

Qlost < 0.5 &-minor discharge Figure L-Predicted discharge current transients.

Qlost < 2 PC-moderate discharge

QloSt < 10 PC-severe discharge

. The current in a discharge pulse can be modeled in anyof several ways, such as approximation by square,triangular, or double exponential pulses or by aresistance-inductance-capacitance (RLC) series circuit.

As an example, an RLC model yields a current given by

zDp)exp(-2LRt)-exP(df)-JxP(-df)

where

&(!L)‘-(&>ln *

The dielectric surface voltage change with time can becomputed from

dVzD=c-$ *

where V” is the value of the dielectric surface voltage attime t. By integrating this expression the charge loss canbe determined. The resistance, inductance, andcapacitance vaues can be adjusted to produce a desiredcharge loss simulating the estimated stored charge that ispredicted in the discharge. The duration of the pulse isthe time required for the current to go to zero. Typicalexamples of this procedure for discharge currents areshown in figure 1 for the cases where the dielectric ischarged to -2000, - 5000, and - 10 000 V just before tidischarge. Figure 2 shows the associated changes indielectric surface voltages.

Dielectric surfacevoltage,

100 200 300Time, ns

. , I / -0 100 200 300Time, ns

Figure Z.--change in dielectric surface voltage due to discharges.

6

2.3.2 Buried charge breakdownsThis section refers to the situation where charges have

sufficient energy to penetrate slightly below the surfaceof a dielectric and are trapped. If the dielectric surface ismaintained near zero potential due to photoelectron orsecondary-electron emission, strong electric fields mayexist in the material.

This can lead to electric fields inside the material largeenough to cause breakdowns. Breakdown can occurwhenever the internal electric field exceeds 2 x 10s V/cm.As an example, figure 3 illustrates the electric fields insidea Teflon film irradiated by a 120keV electron beam. Notethat the field changes sign inside the material, at a depthof about 0.5 pm for this example. The zero field depthdivides the dielectric into two regions for field buildup,labeled regions I and II in the figure. Simple models forthe currents and fields in these regions can be used toobtain estimates of the conductivity required to avoidbuildup of fields larger than 2 x 105 V/cm.

The differential equation relating currents and fields ineach of the two regions (for a linear dielectric in onedimension) is

where c is the dielectric constant, S(X) is the conductivityat depth X, E&t) is the electric field, and J(X) is thecurrent density. The solution to this equation, assumingJ(x) and S(X) are independent of time, is

2xldr 250 Ti me,s.

-

Region II

-6 c-a ’ 1 I 1 I

0 .5 1.0 1.5 20Depth, irm

Figure 3.-Evolution of electric field with time in 0.13.mm (5.mil)Teflon. Calculated for 120keV monoenergetic electron beam at .,0.1 nAbn2.

E&t) = E&)exp [qp]

X (l-exp[l-exp[*]J)

where E&) is the field at t = 0. At long times this reducesto the form E =J/s. Appropriate identification of J foreach region can thus be used to estimate s values.

The J for each region can be identified by consideringthe equivalent circuit diagram of figure 4. Here the nodelabeled 0 represents the zero field point, JF is the currentfrom the front to space, and JB that to spacecraft ground(assumed equivalent to plasma ground). In region I thestrongest electric fields are near the substrate, and theappropriate conductivity is that for the unirradiateddielectric material. The current in this region can be asmuch as 33 percent of the injected current. Substormcurrent densities are typically in the range 0.1 to 1 .OnA/cmZ, giving a value of -0.3 nA/cm2 for JB. Thisyields an estimate of s- 1.5 x lo-15 mho/cm for theminimum allowable dark conductivity.

In region II the largest electric fields will develop nearthe front surface. In this region the conductivity includesradiation-induced terms and is generally higher than thedark conductivity. However, because the currents arealso larger, strong fields can develop. The maximum fieldis JF//s(x= 0), where JF can range from a small initialvalue to a large fraction of the incoming current density.This yields s - 5 x 10 - 15 mho/cm for the minimum totalconductivity.

Note that the form of the internal electric field isdetermined by the spectrum of the incoming electrons, sothat the conductivity guidelines derived here areapproximate. Better estimates of the field underparticular circumstances should be used if available.

berpy 1from Lsubstotm T

II Metal izatfon

JF (region II) grounded toSpaCWilft

A 1%=e= plasma

ground

Figure 4.-Ciicuit with which to estimate internal electric fields.

2.3.3 Spacecraft-tospace breakdowns

Spacecraft-to-space breakdowns are generally similarto dielectric surface breakdowns but involve only smalldischarges. It is assumed that a strong electric field existson the spacecraft surfaces -usually due to i\ gwtwt ricinter facing of metals and diclcct rics. This arrangemctt tperiodically triggers a breakdown of the spacecraft-to-space capacitor. Since this capacitance tends to be of theorder-of 2.x lo- 10 F, these breakd&vn transients shouldbe small and rapid. .

.: , w

2.4.1 Gumped-element modelingLumped-element models (LEM) have been used to

define ihe surface ’ charging response to environmentalf&x& (Robin&n and Holman, 1977; Inouye, 1976;Massaro et al,, 1977; Massaro and Ling, 1979) and arecurrently ‘used to predict interior structural currentsresulting from surface discharges. The basic philosophyof a lumped-element model is that spacecraft surfacesand structures can be treated as electrical circuitelements-resistance, inductance, and capacitance. Thegeometry of the spacecraft is considered only to group orlump areas into nodes within the electrical circuit in muchthe same way as surfaces are treated as nodes in thermalmodeling. Therefore, these models can be made as simpleor *as complex & is considered necessary for thecircumstances.

LEM’s developed to predict surface charging rely onthe use of current input terms applied independently tosurfaces. Since there are no terms relating the influenceof charging by one’ area on the incoming flux to otherareas, the predictions usually result in larger negativevoltages than actually observed. Other modelingtechniques, such as NASCAP, more realistically treatthese three-dimensional effects and predict surfacevoltages closer to those measured.

The LEM’s for discharges assume that the structurecurrent transient is generated by capacitive coupling tothe discharge site and is transmitted in the structure byconduction only. An analog circuit network isconstructed by taking into consideration the structureproperties and. the geometry. This network must considerthe principal current flow paths from the discharge site toexposed conductor areas-the return path to spaceplasma ground. Discharge transients are initiated atregions in this network selected as being probabledischarge sites by surface charging predictions or other

means. Transient characteristics are controlled bychoosing values of resistance, capacitance, andinductance to space. The resulting transients within thet~~twrk cat\ be solved by using network cmpu&[r;tttsicttt circuit rtttutysi+ pt’r~~t~~tttls SUctr its ISl’lCl~ 01Sl’lW2. .

The *proccdurc is illtrst rat cd itt t lrc follow i trg sitttpli t’ictlexample. Consider a three-axis-stabilized spacccraf’t(fig. 5) with a shaded dielectric area adjacent to aconductor. The spacecraft is charged by a substorm suchthat the structure potential is about -2.5 kV while theshaded dielectric is at about - 5900 V. These values wereobtained from NASCAP runs. Figure 6 shows onesection of that spacecraft where a discharge could occur.According to the breakdown criteria given in section2.3.1, a discharge should occur that would eject - 3 @Cof charge in about 0.15 p. .

A very simplified, single-path lumped-element modelto simulate the discharge conditions in a spacecraft isshown in figure 7. It is assumed that the spacecraft ischarged relative to the space plasma potential by asubstorm environment. The spacecraft and dielectric aredifferentially charged to -2500 and -5000 V, respec-tively, at which time a discharge occurs. The dischargemodel assumes that the discharge time is short comparedwith the charging times-when switch Sl closes, S2 isassumed to open. The discharge-pulse-shaping networkallows whatever charge is assumed to be stored in thedielectric to leave in a controlled fashion (fig. 1). Thetransient caused by the discharge is capacitively coupledinto the siructure. The single-path representation of thestructure is modeled as a resistor, capacitor, and inductorchosen to produce an underdamped oscillation with afrequency of about 10 MHz-the estimated value of thestructure resonance.

,

The discharge results in the dielectric surface voltagebecoming more positive. This forces the spacecraftvoltage (relative to space) also to become more positive.Eventually, the spacecraft must return to its substorm-driven value, and this can be estimated by assuming thatthe vehicle is a capacitor being recharged with a giventime constant (fig. 8(a)). Here, the spacecraft potentialrises for 200 ns, at which time it returns rapidly to itsoriginal value of -2500 V.

The current induced in the structure by this discharge isshown in figure 8(b). The first 200 ns corresponds to theresponse to the discharge. The flatter region at 200 nscorresponds to the period. in which the structure isrecharging. The oscillations beyond 200 ns are the ringingcurrent at the structure frequency. This ringing is dampedout by the structure resistance. It should be stressed thatthis is an extremely simplified model used to explain acomplex interaction. In reality, there would be manypaths for current flows from the essentially point sourceof a discharge throughout the structure back to *space.This produces Complex wave patterns in the structure that

8

Figure 5 .-Three-axis-stabilized geosynchronous satellite.

OSR array-, Discharge

Figure 6.-External surface discharge modeling-spacecraft model.

Dielectricsurface Dielectric

RP LPHo - -Pulse shape 4

network Space plasmaground

iIStructure I

csL&g-=&!c*Structureground

Figure 7.-External surface discharge modeling--lumped-elementmodel.

40u r

v)-4 .W I I I I 1

0 200 400 600 800 1000Time, ns,

(a) Predicted change in spacecraft voltage due to discharge.(b) Current induced in structure by discharge.

Figure 8.-Spacecraft response to discharge transient.

are difficult to follow. On the ‘other hand, the simplemodel generates the generalized pattern but implies farlarger currents than actual, where the total current flowsthrough the single circuit. For a real case the completeanalysis must be conducted.

9

The changing current in the structure generates achanging magnetic field, which induces a voltage in anadjacent cable or unit. This is illustrated in figure 9 alongwith the equivalent electrical circuit. The voltage gener-ated in a short cable by the structure current is shown infigure 10. The oscillating pattern is distorted while thedischarge is under way but changes to a damped ringingpattern afterward. The tesults illustrate that voltagesinduced in cables can be significant and can persist afterthe discharge . session. Given the voltage ‘and theelectronic impedance, it is possible to evaluate whether aunit would be susceptible to transient upsets.

2.4.2 Specification andProgram (SEMCAP)

Electromagnetic Compatibility

8 - magnetic fieldA - area formed by structure

. and harness

Figure 9.-Lumped-element model for cable coupling computation.(Assume h > cable radius. Structure current generates magnetic fieldthat induces voltage in cab&; cable responds with its electricalcharacteristics.)

Numerous programs have been developed to study theeffects of electromagnetic coupling on circuits. Suchprograms have been used to compute the effects of anelectromagnetic pulse (EMP) and that of an arcdischarge. One program, the- Specification andElectromagnetic Compatibility Program (SEMCAP)developed by TRW Incorporated (e.g., SEMCAPProgram Description, Ver. 7.4, 1975, Heidebrecht), hassuccessfully analyzed the effects of arc discharges onactual spacecraft -the Voyager series.

SEMCAP was originated to calculate cross couplingfrom source circuits to other circuits in a spacecraft. Arcdischarges were modeled in a manner compatible with theSEMCAP input requirements, and the effects on numer-ous spacecraft circuits were estimated. That process isdescribed.. more fully below.

Briefly., SEMCAP permits modeling the interboxharness cabling and the input and output interfaces foreach box on a spacecraft. The interaction of signals on agiven wire with those on every other wire is computed interms of the physical configuration and terminatingimpedances. By using integration in the frequencydomain over the bandwidths of the coupling networksand the receptor circuits, SEMCAP computes the peakvoltage at each receptor due to each source. The designercan then compare this predicted peak voltage from each

$ -1o

-300

I I I400 600 800

Time. ns

Figure IO.-Voltage generated in cable due to structure currents.

(1) Selection of diagnostic points and stimuluslocation

(2) Prediction of spacecraft circuit responses to teststimuli

(3) Limitation of test stimuli to benign levels(4) Extrapolation of test responses to those expected at

other locations(5) Prediction of spacecraft responses to in-flight arcs

For illustration, the SEMCAP analysis done for Voyageris described in appendix C.

1000

source to the threshold of susceptibility or damage of thereceptor. This process identifies the most troublesomesources and the most susceptible receptors. Seeing these

3.0 Spacecraft Design Guidelinesresults suggests where to modify spacecraft design, if This section contains recommendations on designnecessary. . techniques that should be followed in hardening space-

Roughly 240 generators and 240 receptors can be craft systems to spacecraft charging effects. To minimizemodeled by SEMCAP. The SEMCAP code for arc repetition and to make recommendations as brief asdischarges allows possible, this section is divided into two parts: guidelines

.

10

that are generally applicable, and ideas and commentsthat are more applicable to a particular subsystem, suchas the power subsystem. It is suggested that all readersreview the general guidelines section and read componentand subsystem sections for areas of specific concern.

3.1.1.3 wiring and cable shields.-All wiring andcabling exiting the shielded “Faraday cage” portion ofthe spacecraft (section 3.1.3) should be shielded. Thosecable shields and any other cable shields used for ESDpurposes shall be bonded to the Faraday cage at the entryto the shielded region as follows: .

3.1 General Guidelines3.1.1 Grounding

(I) The shield shall be terminated 360” around a metalshielded backshell, which is in turn terminated to thechassis 360” around the cabling.

All conducting elements, surface and interior, shouldbe tied to a common electrical ground, either directly or

(2) The shield ground shall not be terminated by using

through a charge bleedoff resistor.a pin that penetrates the Faraday cage and receives itsground inside the shielded region.

3. I. I. I Structure and mechanical parts. -Allstructural and mechanical parts, electronics boxes,enclosures, etc., of the spacecraft shall be electricallybonded to each other. All principal structural elementsshall be bonded by methods that assure a direct-current(dc) resistance of less than 2.5 mQ at each joint. Thecollection of electrically bonded structural elements isreferred to as “structure” or structure ground. Theobjective is to provide a low-impedance path for anyESD-caused currents that may occur and to provide anexcellent ground for all other parts of the spacecraftneeding grounding. If structure ground must be carriedacross an articulating joint or hinge, a ground strap, asshort as possible, shall carry the ground across the joint.Relying on bearings to serve as a ground path isunacceptable. If structural ground must be carried acrosssliprings on a rotating joint, at least two, and preferablymore, sliprings shall be dedicated to the structural groundpath, some at each end of the slipring set. The bond tostructure shall be achieved within 15 cm of the slipring oneach end . of the rotating joint. Sliprings chosen forgrounding should be away from any sliprings carryingsensitive signals.

3.1.1.2 Surface materials.-All spacecraft surface(visible, exterior) materials should be conductive in anESD sense (section 3.1.2). All such surface materials shallbe electrically bonded (grounded) to the spacecraftstructure. Because they are intended to drain spacecharging currents only, the bonding requirements are lesssevere than those for structural bonding. The dcimpedance to structure should be compatible with thesurface resistivity requirements: that is, less than about10W from a surface to structure. The dc impedance mustremain less than 109 Q over the service life of the bond invacuum, under temperature, under mechanical stress,etc.

(3) A mechanism shall be devised that automaticallybonds the shield to the enclosure/structure.ground at theconnector location, or a ground lug that uses less than 15cm of ground wire shall be provided for the shield andprocedures that verify that the shield is grounded at eachconnector mating shall be established.The other end of the cable shield shall be terminpted inthe same manner. The goal is to maintain shieldingintegrity even when -some electronics units must belocated outside the basic . shielded region of thespacecraft.

3.1.1.4 Electrical and electronic grounds. -Signaland power grounds require special attention in the waythey are connected to the spacecraft structure ground. .For ESD purposes a d i r ec t w i r ing o f a l lelectrical/electronics units to structure is most desirable.In particular, one should not have separate ground wiresfrom unit to unit or from each unit to a single point pnthe structure.

If the electronic circuitry cannot be isolated frompower ground, signal ground may be referenced tostructure with a large (> 10 kQ) resistor. Once again, box-to-box signals must be isolated to prevent ground loops.This approach must be analyzed to assure that it isacceptable from an ESD standpoint.

In some cases it is necessary to run signal and powerground lines in harnesses with other space vehicle wiring.This should be avoided where possible and limited whereconsidered necessary. Excessively long runs of signalground lines should be eliminated.

3.1.2 Exterior surface materialsFor differential charging control, all spacecraft

exterior surfaces shall be at IeaSt partially conductive.The best way to avoid differential charging of spacecraftsurfaces is to make all surfaces conductive and grounded

11

to the spacecraft sfrucfure. However, typical spacecraftsurface materials often include insulating films such asMylar, Kapton, Teflon,. fiberglass, glass, quartz, or otherdielectric materials. It should be recognized in the designphase that there may be areas for which use of conductive.surfaces is particularly crucial, such as areas adjacent toreceivers/antennas operating at less than 1 GHz, sensitivedetectors’(Sun and Earth dktectors, etc.) or areas wherematerial contaniinatioti or thermal control is critical. Forthese applications use of indium tin oxide (ITO) coatingsis recommended. *

(1) Isolated conductors must be grounded with lessthan 106 0 to structure. .

(2) Materials applied over a conductive substrate musthave bulk resistivities of less than 1011 Q-cm. .

(3) Materials applied over a dielectric area must begrounded at the edges and must have a resistivity lessthan 109 “ohms per square.“1These requirements are more strict than the precedingrelations, which include effects of spacecraft geometry.

In all cases the usage or application process must beverified by measuring resistance from any point on thematerial surface to structure. Problems can occur. Forexample, one case was observed where a nonconductiveprimer was applied underneath a conductive paint; thepaint’s conductivity was useless over the insulatingprimer.

This section first defiries the conductivity requirementsfor spacecraft surface materials. Materials that aretypically used are then evaluated axid their usage isdiscussed. Analysis ‘is stiggested to estimate the effects of. !any dielectric stirfaces that may remain on the spacecraft.At the.bo@asion. of this section, use of materials with ahi&h secondary &ctron yield is discussed.. .

3.1.2. I Surface conductivity requirements.-Todischarge surfaces that are being charged by spaceplasmas, a high resistivity to grourid can be toleratedbecause the plasma charging currents are small. Thefollowing guidelin& are recommended:

All grounding methods must be demonstrated to beacceptable over the service life of the spacecraft. It isrecommended that all joint resistances and surfaceresistivities be measured to verify compliance with theseguidelines. Test voltages should be at least 500 V.Grounding methods must be able to handle currentbleedoff from ESD events, vacuum exposure, thermalexpansion and contraction, etc. As an example, paintingaround a zero-radius edge or at a seam between twodissimilar materials could lead to cracking and a loss ofelectrical continuity at that location.

3.1.2.2 Surface materials.-By the proper choice ofavailable materials the differential charging of spacecraftsurfaces can be minimized. At present, the only provenway to eliminate spacecraft potential variations is bymaking all surfaces conductive and tying them to acommon ground.

Surface coatings in use for this purpose includeconductive conversion coatings on metals, conductivepaints, and transparent partially metallic vacuum-deposited films, such as indium tin oxide. Table IIdescribes some of the more common acceptable surfacecoatings and materials with a successful use history.Table III describes other common surface coatings andmaterials that should be avoided if possible.

The following materials have been used to provideconducting surfaces on the spacecraft:

(I) Conductive materials (e.g., metals) must begrounded to-. structure with the smallest resistancepossible

Rc lO?/A, Q.

where A is the exposed surface area of the conductor insquare centimeters.

-(2) Partially conductive surfaces (e.g., paints) appliedover .a conductive substrate must have a resistivity-thickness product

rtS2X 109, Q-cm2 .

where r is the material resistivity in ohm-centimeters and Iis the material thickness in centimeters.

(3) Partially conductive surfaces applied over adielectric and grounded at the edges must have materialresistivity such thatrh2IS4x 109, &cm2

where r and t are as above and h is the greatest distanceon a surface to a ground point, in centimeters.

These guidelines depend on the particular geometryand aiplication. A simplified set of guidelines is supplied

f “Ohms per square” is defined as the resistance of a flat sheet of thematerial measured from one edge of a square section to the oppositeedge. It can be seen that the size of the square has no effect on the

for early design activities: numeric value.

li

TABLE II. - SURFACE COATINGS AND MATERIALS ACCEPTABLE FOR SPACECRAFT USE

Material

Paint(carbon black)

GSFC NS43"paint (yellow)

Indium tin oxide(250 nm)

Zinc ortho-titanate paint(white)

Alodyne

Possibly the most conductive white paint; adhesion diffi-cult without careful attention to application procedures

Conductive conversion coatings of magnesium, aluminum, etc.,are acceptable

"GSFC denotes Goddard Space Flight Center. M

TABLE III. - SURFACE COATINGS AND MATERIALS TO BE AVOIDED FOR SPACECRAFT USE7

Material Comments --Anodyze Anodyring produces a high-resistivity surface; to be

avoided. The surface is thin and might be acceptable ifanalysis shows stored energy is small

Fiberglassmaterial

Resistivity is too high

Paint (white) In general, unless a white paint is measured to beacceptable, it is unacceptable

Mylar (uncoated) Resistivity is too high

Teflon . Resistivity is too high. Teflon has a demonstrated long-time(uncoated) charge storage ability and causes catastrophic discharges

Kapton Generally unacceptable, due to high resistivity. However,(uncoated) in continuous-sunlight applications if less than 0.13 mm

(5 mils) thick, Kapton is sufficiently photoconductive foruse

Silica cloth Has been used as antenna radome. It is a dielectric, butbecause of numerous fibers, or if used with embeddedconductive materials, ESD sparks may be individually small

Quartz and It is recognized that solar cell coverslides and second-glass surfaces surface mirrors have no substitutes that are ESD .

acceptable. Their use must be analyzed and ESD testsperformed to determine their effect on neighboringelectronics

(1) Vacuum-metalized dielectric materials in the form carbon-filled Teflon, or carbon-filled polyester onof sheets, strips, or tiles. The metal-on-substrate combi- Kapton (Sheldahl black Kapton)n&ions include aluminum, gold, silver, and Inconel on (4 j Conductive adhesivesKapton, Teflon, Mylar, and fused silica. (5) Exposed conductive facesheet materials (graphite/

(2) Thin, conductive front-surface coatings, especially epoxy or metal)indium tin oxide on fused silica, Kapton, Teflon, or (6) Etched metal grids or bonded (or heat embedded)dielectric stacks metal meshes on nonconductive substrates

(3) Conductive paints, fog (thin paint coating), (7) Aluminum foil or metalized plastic film tapes

13

Because of the variety in the configuration and propertiesof these materials, there is a corresponding variety in theapplicable grounding techniques and specific concernsthat must be addressed to’ insure reliable in-flightperformance.

The following practices have been found useful:

(1) Conductive adhesives should be used to bond fusedsilica, Kapton, and Teflon second-surface mirrors toconductive substrates that are grounded to structure. Ifthe substrate is not conductive, metal foil or wire groundlinks should be laminated in the adhesive and bolted tostructure. Only optical solar reflectors (OSR’s) withconductive (Inconel) back surfaces should be used. *

(2) When conductive adhesives are used, the long-termstability of the materials system must be verified, partic-ularly conductivity in vacuum after thermal cycling,compatibility of the materials (especially for epoxyadhesive) in differential thermal expansion, and long-term resistance to galvanic corrosion.

. (3) Metalized Teflon is particularly susceptible toelectrostatic discharge degradation, even when grounded.Avoid using it. If there is no substitute for a specificapplication, the effects of electromagnetic interference(EMI), contamination, and optical and mechanicaldegradation must be evaluated.

(4) Paints should be applied to grounded, conductivesubstrates. If this is not possible, their coverage should beextended to overlap grounded conductors.

(5) Ground tabs must be provided for free-standing(not bonded down) dielectric films with conductivesurfaces.

(6) Meshes that are simply stretched over dielectricsurfaces are not effective; they must be bonded or heatsealed in a manner that will not degrade or contaminatethe surface.

(7) There are several techniques for grounding thin,conductive front-surface coatings such as indium tinoxide, but the methods are costly and have questionablereliability. The methods include welding of ground wiresto front-surface metal welding contacts, front-surfacebonding of coiled ground wires (to allow for differentialthermal expansion) by using a conductive adhesive, andchamfering the edges of OSR’s before IT0 coating topermit contact between the coating and the conductiveadhesive used to bond the OSR to its substrate.

Grounding techniques for OSR’s include chamferingedges and bonding with conductive adhesive and front-surface bonding or welding of ground wires. Bondingdown solar cell covers with conductive adhesive is notapplicable. For multilayer insulation (MLI), extendingthe aluminum foil tab to the front surface is suitable.

3.2.2.3 Nonconductive surfaces.-If the spacecraftsurface cannot be made 100 percent conductive, ananalysis must be performed to show that the design isacceptable from an ESD standpoint. Note that not all

dielectric materials have the same charging or ESDcharacteristics. The choice of dielectric materials can sig-nificantly affect surface voltage profiles. For example, ithas been shown (Bever, 1981) that cerium-doped micro-sheet charges to much lower potentials under electronirradiation thm fused silica, ~lnd it t hert’ti~rr tnrrv 11~preferred as a solar array coverslide trratcrial.

An adequate analysis preceding the selection ofmaterials must include spacecraft analysis to determinesurface potentials and voltage gradients, spark dischargeparameters (amplitude, duration, frequency content),and EM1 coupling. The cost and weight involved inproviding adequate protection (by shielding and electricalredesign) could tilt the balance of the trade-off to favorthe selection of the newer, seemingly less reliable(optically) charge control materials that are more reliablefrom spacecraft charging, discharging, and. electro-magnetic interference points of view.

The “proven” materials have their own cost, weight,availability, variability, and fabrication effects. In addi-tion, uncertainties relating to spacecraft charging effectsmust be given adequate consideration. Flight data haveshown apparent optical degradation of standard, stablethermal control materials (e.g., optical solar reflectorsand Teflon second-surface mirrors) that is far in excess ofground test predictions, part of which could be the resultof charge-enhanced attraction of charged contaminants.In addition, certain spacecraft anomalies and failuresmay have been reduced or avoided by using chargecontrol materials.

.

Ironically, after an extensive effort to have nearly all ofthe spacecraft surface conductive, the remaining small

’patches of dielectric may charge to a greater differentialpotential than a larger area of dielectric would. On theshadowed side of a spacecraft, a small section ofdielectric may be charged rapidly while the bulk of thespacecraft remains near zero potential because ofphotoemission from sunlit areas.

A spacecraft with larger portions of dielectric mayhave retarding electric fields because the dielectricdiminishes the effects of the photoemission process. As aresult, the spacecraft structure potential may go morenegative and thus reduce the differential voltage betweenthe dielectric and the space&aft.

The lesson to be learned is that all dielectrics must beexamined for their differential charging. Each dielectricregion must be assessed for its breakdown voltage, itsability to store energy, and the effects it can have onneighboring electronics (disruption or damage) and

.

surfaces (erosion or contamination).3.1.2.4 Surface secondary emission ratios.-Other

means to reduce surface charging exist but are not welldeveloped and are not in common usage. One suggestionfor metallic surfaces is an oxide coating with a highsecondary electron yield. This concept, in a NASCAPcomputer program simulation, reduced the absolute

14

charging of a spacecraft dramatically and reduceddifferential charging of shaded Kapton slightly. Anyselected materials should be carefully analyzed to insurethat they do not create problems of their own and thatthey work as intended over their service lives.

Another concept to reduce charging, the neutralplasma beam, is discussed in section 5.0.

3.1.3 ShieldingThe primary spacecraft structure, electronic

component enclosures, and electrical cable shields shallprovide a physically and electrically continuous shieldedsurface around all electronics and wiring (Faraday cage).The primary spacecraft structure should be designed asan electromagnetic-interference-tight shielding enclosure(Faraday cage). The purposes of the shielding are (1) toprevent entry of space plasma into the spacecraft interiorand (2) to shield the interior electronics from the radiatednoise of an electrical discharge on the exterior of thespacecraft. All shielding should provide at least 40dBattenuation of radiated electromagnetic fields associatedwith surface discharges. An approximately l-mm thick-ness of aluminum or magnesium will generally providethe desired attenuation. This enclosure should be as freefrdm holes and penetrations as possible. Manypenetrations can be made relatively electromagneticinterference tight by use of well-grounded metallicmeshes and plates. All openings, apertures, and slits shallbe eliminated to maintain the integrity of the Faradaycage.

The metalization on multilayer insulation isinsufficient to provide adequate shielding. Layers ofaluminum foil mounted to the interior surface andproperly grounded can be used to increase the shieldingeffectiveness of blankets or films. Aluminum honeycombstructures and aluminum facesheets can also provide

* significant attenuation. Electronic enclosures andelectrical cables exterior to the main Faraday cage regionshould also be shielded to extend the coverage of theshielded region to 100 percent of the electronics.

Cable shields exterior to the Faraday cage shall main-tain and extend the cage region from their exit/extranceof the main body of the spacecraft. Cable shields shouldbe fabricated from aluminum or copper foil, sheet, ortape. Standard coaxial shielding or metalized plastic tapewraps on wires do not provide adequate shieldingprotection and should not be used. Shields shall beterminated when they enter the spacecraft structure fromthe outside and carefully grounded at the entry point.Braid shields on wires should be soldered to any overallshield wrap and grounded at the entrances to thespacecraft. Conventional shield grounding through aconnector pin to a spacecraft interior location should notbe used.

Electrical terminators, connectors, feedthroughs, and

externally mounted components (diodes, etc.) should beelectrically shielded and all shield caps tied to thecommon structural ground system of the space vehicle..

3.1.4 FilteringElectrical filtering should be used to protect circuits

from discharge-induced upsets. All circuits routed intothe Faraday cage region, even though their wiring is inshielded cabling, run a higher risk of having ESD-causedtransient voltages on them. Initial design planning shouldinclude ESD protection for these circuits. It is recom-mended that filtering be applied to these circuits unless *analysis shows that it is not needed.

The usual criterion suggested for filtering is to elimi-nate noise below a specific time duration (i.e., above aspecific frequency). On the Communications TechnologySatellite (CTS), in-line transmitters and receivers wereused that effectively eliminated noise pulses of less than5.~ duration. Similar filtering concepts might include avoltage threshold or energy threshold. Filtering isbelieved to be an effective means of preventing circuitdisruption and should be included in system designs. Anychosen filtering method should have analyies and tests tovalidate the selected criteria; Filters should be rated towithstand the peak transient voltages over the mission.life.

3.1.5 ProceduresProper handling, assembly, inspection, and test

procedures shall be instituted to insure the electricalcontinuity of the space vehicle grounding system. Thecontinuity of the space vehicle electrical groundingsystem is of great importance ‘to the overall designsusceptibility to spacecraft charging effects. In addition itwill strongly affect the integrity of the space vehicleelectromagnetic capability (EMC) design. Properhandling and assembiy procedures must be followedduring fabrication of the electrical grounding system. Allground ties should be carefully inspected and dcresistance levels should be tested during fabrication andagain before delivery of the space vehicle. A final checkof the ground system continuity during preparation forspace vehicle launch is desirable.

3.2 Subsystem GuidelineqThe guidelines in this section are divided by spacecraft

subsystem. Designers of specific subsystems should readthe applicable portions of this section and, in addition,review the general guidelines (table IV).

3.2.X ElectronicsThe general guidelines apply.

3.2.2 Power systemsSee table IV. In addition, the follo\ling specific

guidelines apply. .3.2.2.1 Solar panel groundbg.-Solar army panels

and substrates shall be electrically grounded to thestructure. Solar-array panels and conductive sections ofsubstrates’and honeycomb.-ihould be grounded to eachother with &ou&hg jumpers and the entire networkgrounded to the space vehicle stru&re with less than2$&k .dcy re;sista&q per jgint; Deployable panels onthree&is-stabilized vehicles can be. grounded to thestructure .through $iprings where necessafy. A groundwire can be-used to bond together each lateral strip orrow of solar cells.. :

’ 3,2.2.:2. I c, .. * .

&w panel fabrication.4olar arny panelsshah t&e materials and fabrication - techniqk tomini&ze.flectrostatic dischkge effects. Solar panel backsurfaces, edges, and. honeycomb should be. groundedconductors. Conductive black’ paint is suitable for therear s&ace .of the solar panel. Solar panel edges can bewyal$e&with .-grounded- conductive tape. The frontsurface. of *the solar array * co&i&s of nonconductivecoversli.des a& -gaps so&times potted with noncon-ductive.adhesive’ for electrical ,desigxi reasons. The p.ottingthickness should be ‘the minimum required. The front. .surfaces of coverslides may be coated with a conductive,transparent coating of grounded indium tin oxide ifrequired. Such coatings typically reduce transmission by5 to 10 percent and are generally used when absolutecharging must be controlled.

3.2.2.3 PO wer system electrical design. -Powersystem electrical’ design shall incorporate features toprotect against transients due to electrical discharge.Spark discharges from solar arrays should be anticipated,and the electrical design- of the power system mustprovide adequate protection. The following designpractices will help in reducing the effects of such sparkdischarges.

(1) Clamp solar array wiring, preferably at the entry tothe spacecraft Faraday cage, but definitely before itenters the power supply. .

(2) If solar array wiring is not clamped at the entrypoint to the Faraday cage, shield the wiring from thatpoint to the power supply.

(3) Use solar array diodes with forward current ratingsthat anticipate expected ESD transient currents.

(4) Perform analysis and testing to verify the powersystem electrical design for survivability or immunity to. . . .spacecraft charging effects. - .. . . .

.

3.2.3 l&ha&al and structural. .

See table IV. In addition, the following specific guide-line applies: Conductive honeycomb and facesheets shallbe electrically grounded to the structure.

Aluminum . honeycomb substructures require specialconsideration for electrical grounding. Techniques forgrounding conductive honeycomb and facesheets include l *rivets, copper wires, and metal inserts. *’ ’.

Care should be taken to establishground ties at severallocations on the honeycomb structure and to maintain..ground continuity through all honeycomb parts and.facesheets. For example, a recommended method ofusing copper wires involves sewing the wires transverselyat shallow inclination angles through the honeycomb(making contact with several of the cell walls). The wiresshould be installed at maximum intervals of 30 cm acrossthe structure. Ground wires should then be bolted to thestructure. Electrical inspection of grounding interfacesfor honeycomb structures applies. .

3.2.4 Thermal controlSee table IV. In addition, the following specific guide:

lines apply: .3.2.4.1 Thermal blankets.-All metalized surfaces @

multilayer insulation (MLI) blankets shall be electrically

TABLE IV. - SUBSYSTEM GUIDELINES - APPLICABLE SECTIONS,

Subsystem and design Applicable sectionstechnology

3.1.1 '3.1.2 3.1.3 3.1.4 3.1.5 Extra ..

Electronics

Power '. ..

Mechanical andstructure

X X X X X

X * x X X X 3.2.2

X X x . 3.2.3 *

Thermal X

Radiofrequency ' Xand communications

X X 3.2.4..

X X X X 3.2.5

Attitude control X X X X X 3.2.6

Payloads . X X X 3 X 3.2.7 I1

IQ

grounded to the structure. The metalized multilayersurfaces should be electrically grounded to each otherby ground tabs at the blanket edges. Each tab shouldbe made from a 2.5-cm-wide strip of O.OOS-cm-thickaluminum foil. The strip should be accordion folded andinterleaved between the blanket layers to give a 2.5 by2.5-cm contact area with all metalized surfaces and theblanket front and back surfaces. Nonconductive spaceror mesh material must be removed from the vicinity ofthe interleaved tab. The assembly should be held in placewith a metallic nut and bolt that penetrates all blanketlayers and captures 2.0-cm-diameter metallic washerspositioned on the blanket front and back surfaces andcentered in the 2.5. by 2.5,cm tab area. The washers mayhave different diameters, with the inner surface of thesmaller washer recessed to insure maximum peripheralcontact area between the interleaved foil strip and eachmetalized blanket surface. The tab should be grounded tostructure by a proven technique such as a wire that is asshort as possible (15 cm maximum) or conductive Velcro.

Redundant grounding tabs on all blankets are requiredas a minimum. Tabs should be located on blanket edgesand spaced to minimize the maximum distance from anypoint on the blanket to the nearest tab. Extra tabs may beneeded on odd-shaped blankets to meet one additionalcondition: any point on a blanket should be within 1 m ofa ground tab.

The following practices should be observed duringblanket design, fabrication, handling, installation, andinspection:

(1) Verify layer-to-layer blanket grounding duringfabrication.

(2) After installation, verify less than IO-Q dcresistance between blanket and structure.

(3) Close blanket edges (cover, fold in, or tape) toprevent direct irradiation of inner layers.

(4) Do not use crinkled, wrinkled, or creased metalizedfilm material.

.

(5) Handle blankets carefully to avoid creasing of thefilm or possible degradation of the ground tabs.

(6) If the blanket exterior is conductive (paint, indiumtin oxide, “fog”), make sure that it contacts the groundtab.

3.2.4.2 Tbmal control Zouvem-Ground the bladesof thermal control louvers. A fine wire with minimaltorque behavior or a fine slip brush can do the job withacceptable torque constraints.

3.2.5 Communications systems

See table IV. In addition, the following specific guide-lines apply.

3.2.5.1 Antenna grounding.-Antenna elementsshall be electrically grounded to the structure.Implementation of antenna grounding will require

careful consideration in the initial design phase. All metalsurfaces, booms, covers, and feeds should be groundedto the structure by wires and metallic screws (dc shortdesign). All waveguide elements should be electricallybonded together with spotwelded connectors andgrounded to the spacecraft structure. These elementsmust be grounded to the Faraday cage at their entrypoints. Conductive epoxy can be used where necessary,but dc resistance of about 1 St must be verified bymeasurements.

3.2.5.2 Antenna apertures.-Spacecraft rf aqtennaaperture covers shall be ESP condictive and grounded.Charging and arcing of dielectric antenna dish surfacesand radomes can be prevented by covering them withgrounded ESD-conductive material. Antenna per-formance should be verified with the ESD coveringinstalled.

3.2.5.3 Antenna reflector surfaces.- Grounded,conductive spacecraft charge control materials shall beused on antenna reflector rear surfaces. Appropriatesurface covering techniques must be selected. Applicablemethods include conductive meshes bonded to dielectricmaterials, silica cloth, conductive paints, or non:conductive (but charge bleeding) paints overlappinggrounded conductors.

3.2.5.4 Transmitters and receivers. -Spacecrafttransmitters and receivers (command line and data line)shall be immune to transients produced by electrostaticdischarge. Transmitter and receiver electrical design mustbe compatible with the results of spacecraft chargingeffects. The EM1 environment produced by spacecraftelectrostatic discharge should be addressed early in thedesign phase to permit effective electrical design forimmunity to this environment. The transmitter, receiver,and antenna system should be tested for immunity toESD’s near the antenna feed. The repetition rate shall beselected to be consistent with estimated arc rates ofnearby materials.

3.2.6 Attitude controlAttitude control electronics packages should be

insensitive to ESD transients. See table IV. Attitudecontrol systems often require sensors that are remotefrom electronics packages for Faraday shielding. Thispresents the risk that ESD transients will be picked upand conducted into electronics. Particular care must betaken to insure immunity to ESD upset in such cases.

3.2.7 Payloads

See table IV. In addition, the following specificguidelines apply.

3.2.7.2 Deployed packages. -Deployed packagesshall be grounded by using a flat ground strap extendingthe length of the boom to the vehicle structure. Severalspacecraft designs incorporate dielectric booms to deploy

17

payloads. The payload electrical system may still require (3) Have a design qualification test sequence mat isa common ground reference, or the experiment may extensive: test all units of hardware, use long tatrequire a link to some electric potential reference. Inthese cases it is recommended that a flat ground strap be

durations, examine many equipment operating modes,apply the environment to all surfaces of the test unit. 1

used to carry this ground tie to the vehicle structure.Electrical wiring extending from the deployed payload tothe spacecraft interior must be carried inside or along thedielectric booms. This wiring should be shielded and theshield grounded at the package end and at the Faradaycageeentrmce. .: ’ I +’ . .

.3;2.7,2 . Ungrounded materials.-Specific items thatcannot be grounded because of system requirements shallundergo tinalysis to assure specified performance in thespacecraft charging environment. Certain space vehiclesmay contain specific items or materials that must not begrounded. For example, a particular experiment mayhave a metallic grid or conducting plate that must be leftungrounded; If small, these items may present nounusual spacecraft charging problems; however, this.should be verified through analysis.

3.2.7.3 DMberate swface potentiuk--If a surfaceon the spacecraft must be charged (detectors on a scienceinstrument, for example), it shail be recessed or shieldedso that the perturbance in surface electrostatic potentialsis less than 10 V. Scientific instruments with the need forexposed surface voltages for measurement purposes, suchas Faraday cups, require special attention to insure thatthe electrostatic. fields they create will not disruptadjacent surface charging or cause discharges by theiroperation. They can be recessed so their fields at thespacecraft surface are minimal or shielded with groundedgrids. An analysis may be necessary to insure that theirpresence is tolerable from a spacecraft chargingstandpoint.

(4) Have a flight hardware test sequence of moremodest scope: delete some units from test if qualificationtests show great design margins, use shorter test dura-tions, use only key equipment operating modes, andapply the environment to a limited number of surfaces.

Ideally, both prototype and flight spacecraft should betested in a charging simulation facility. They should beelectrically isolated from ground and bombarded withelectron, ion, and extreme ultraviolet radiation levelscorresponding to substorm environment conditions.Systems should operate without upset throughout thistest.

Because of the difficulty of simulating the actualenvironment. (space vacuum and plasma parametersincluding species such as ions, electrons, and heavierions; mean energy; energy spectrum; and direction),spacecraft charging tests usually take the form ofassessing unit immunity to electrical discharge transients..The appropriate discharge sources are based on separateestimates of discharge parmeters.

Tests in a room ambient environment employingradiated and injected transients are more convenient..However, these ground tests cannot simulate all effects ofthe real environment because the transient source maynot be in the same location as the region that maydischarge and because a spark in air has a slower risetimethan a vacuum arc. The sparking device’s location andpulse shape must be analyzed to provide the best possiblesimulation of coupling to electronic circuits. To accountfor the difference in risetime, the peak voltage might beincreased to simulate the dI?/dt parameter of a vacuumarc. Alternatively the voltage induced during a test couldbe measured and the in-flight noise extrapolated from the

4.0 Spacecraft Test Techniques measured data.

Spacecraft and systems should be subjected to tran-sient upset tests to verify immunity. It is the philosophyin this document that testing is an essential ingredient in asound spacecraft charging protection program. In thissection the philosophy and methods of testing spacecraftand spacecraft systems are reviewed.

A proper risk assessment will involve a well-plannedtest, predictions of voltage stress ‘levels at key spacecraftcomponents, verification of these predictions during test,checkout of the spacecraft after test, and collaborationwith all project elements to coordinate and assess the riskfactors.

.4.1 Test Phkosophy

The philosophy of the BSD test is identicalthe normal environmental qualification test:.

to that of

(1) Subject the spacecraft to an environmentrepresentative of-that expected.

(2) Make. the environment applied to -the spacecraftmore severe than expected as a safety margin to giveconfidence that the flight spacecraft will survive the realenvironment.

4.2 Simulation of ParametersBecause ESD test techniques are not well established, it

is important to understand the various parameters thatmust be simulated, at a minimum, to perform ’ anadequate test. On the basis of their possibility ofinterference to the spacecraft, the following items shouldbe considered in designing tests: * .

(1) Spark location(2) Radiated fields or structure currents .

(3) Area, thickness, and dielectric strength of thematerial

(4) Total charge involved in the event - *(5) Breakdown voltage l ” ’ ” ” -(6) Current waveform: risetime, width, falltime, and

rate of rise (in amperes per second) *(7) Voltage waveform: risetime, width, falltime, and

rate of rise (in amperes per second)Table V shows typical values as calculated on some

spacecraft. The values listed in this table were compiledfrom a variety of sources, mostly associated with theVoyager and Galileo spacecraft. The values for each item(e.g., those for the dielectric plate) have been assembledfrom the best available information and made into amore or less self-consistent set of numbers. The process isdescribed in the footnotes to table V. See thebibliography for further description and discussion. .

4.3 General Test Methods4.3.1 ESD-generating equipment

Several representative types of test equipment aredescribed in table VI. Where possible, typicalparameters for that type of test are listed.

4.3. I.2 MZL-STD-Is41 arc source.-The MilitaryStandard 1541 (MIL-STD-1541) arc source is commonlyused. The schematic and usage instructions extractedfrom MIL-STD-1541 are presented here as figure Il.

The arc source can be manufactured relatively easily andcan provide some of the parameters necessary to simulatea space-caused ESD event. The only adjustable param-eter for the MIL-STD.1541 arc source is the dischargevoltage (achieved by adjusting the discharge gap and, ifnecessary, the dc supply to the discharge capacitor). As aresult, peak current and energy vary with the discharge .voltages. Since the risetime, pulse width, and falltime aremore or less constant, the voltage and current rates of riseand fall are not independent parameters. This permitssome degree of flexibility in planning tests but notenough to cover all circumstances.

4.3.1.2 Flat-plate capacitor.-A flat-plate capacitorcan be used in several circumstances. Examples ofspacecraft areas that can be simulated by a flat-platecapacitor are (1) thermal blanket areas, (2) dielectricareas such as calibration targets, and (3) dielectric areassuch as nonconductive paints. The chief value of a flat-plate capacitor is to permit a widespread discharge tosimulate the physical path of current flow. This can be ofsignificance where cabling or circuitry is near the area inquestion. Also, ‘the larger size of the capacitor platesallows them to act as an antenna during discharge andthus produce significant radiated fields.

Table VI shows one example of the use of a flat-platecapacitor. Several parameters can be varied, chiefly thearea and the dielectric thickness; both of these affect thecapacitance, the discharge current, and the energy. The

TABLE V. - EXAMPLES OF ESTIMATED SPACE-GENERATED ESD SPARK PARAMETERS

ESD generator

Dielectricplate. toconductive .substrate

Exposedconnectordielectric

Paint on high-gain antenna

Conversioncoating onmetal plate

Paint onoptics hood

Zapac-tance, aLnF

Weakdow!roltage,

2 -

1

* 5

1

1

.36Q.

2.25 16:

ischargecurrentisetime,et *k:

lischarge:urrentpulsesidth,ft 9'n P

lo

15

‘2400

285

600

aComputed from surface.area,.dielectric thickness, and dielectric constant.bComputed from dielectric$hicknesi and material breakdown strength.CComputed from E = l/2 CV :dEstimated based on measured data; extrapolation based on square root of area.eMeasured and deduced from test data.fTo balance total charge on capacitor.gThis was replacement current in longer ground wire; charge is not balanced.

19

TABLE VI. - EXAMPLES OF SEVERAL ESD GENERATORS_ _ - - -_- _ _.____. ___

ES0 generator Capac- Breakdownftance,a voltage,

Energy, Peak OischargeE,

Dischargecurrent,

. c,current current

v 9. nf k&

mJ 1 kgA

risetime, pulsetrB width,ns t 9

n!

MIL-STD-1541 0.035 19 6 80 5 20(auto coil)a

Flat plate 14 5 180 80 35 88020 cm x 20 cmat 5 kV, 0.08 mm

' (3 mil) Mylarinsulation

Flat plate withlumped-elementcapacitor

550 ,450 55 i s 15 (b)

Capacitor direct 1.1 . 320 .056 . 1 3x109 0. injection 10x104

20.

Capacitor arcdischarge

60 1.4 59 1000 ( C 1 801

aparameters were measured on one unit similar to the MIL-STD-1541 design.bRC time constant decay.CValue uncertain.

i

Turns ratioof loo

.

0.1 wIi

it . I 3I

Silicon-I 5opF

4 controlled$ distr ibuted ” Adjus’Me

rectifier rated , I capacitance Wp

at=5AIII

++Carbonelectrodes

Generator must becapable of drivingrelay; a rate of1 pulsepersecondshall be used

Gap mounted on a phenolicboardwith electrodesonadjustableTeflon shafts

Typicalgap-spacingandvoltagebreakdown level

Figure 1 l.-Schematic diagram of MIL-STD-1541 arc source.

discharge voltage of the flate plate can be controlled byusing a needlepoint discharge gap at its edge that iscalibrated to break down before the dielectric. This gapalso affects discharge energy. In this manner, severalmechanical parameters can be designed to yield dischargeparameters more closely taiiored to those expected inspace.

The difficulties of this method include the following:

(1) The test capacitor is usually not as close to theinterior cabling as the area it is intended to simulate (e.g.,it cannot be placed as close as the paint thickness).

(2) The capacitance of the test capacitor may be lessthan that of the area it is intended to simulate. To avoiduncontrolled dielectric breakdown in the test capacitor,its dielectric may have to be thicker than the region itsimulates. If so, the capacitance will be reduced. The areaof the test capacitor can be increased to compensate, butthen the size and shape will be less realistic.

4.3.1.3 Lumped-element capacitors.-Use oflumped-element capacitors can overcome some of theobjections raised about flat-plate capacitors. They canhave large capacitances in smaller areas and thussupplement a flat-plate capacitor if it alone is notadequate. The deficiencies of lumped-element capacitorsare as follows:

(1) They generally do not have the higher breakdownvoltages (greater than 5 kV) needed for ESD tests.

(2) Some capacitors have a high internal resistance andcannot provide the fast risetimes and peak currentsneeded to simulate ESD events.Generally, the lumped capacitor discharge would be usedmost often in lowei voltage applications (to simulatepainted or anodized surface breakdown voltages) and inconjunction with the flat-plate capacitors.

4.3.1.4 Other source equipment.-Wilkenfeld et al.(1982) describes several other similar types of ESDsimulators. It is a useful document if further descriptionsof ESD testing are desired.

4.3. I. 5 Switches. -A wide variety of switches can beused to initiate the arc discharge. At low voltages,semiconductor switches can be used. The MIL-STD-1541 arc source uses an SCR to initiate the spark activityon the primary of a step-up transformer; the high voltageoccurs at an air spark gap on the transformer’ssecondary. Also at low voltages, mechanical switches canbe used (e.g., to discharge modest-voltage capacitors).The problem with mechanical switches is their “bounce”in the early milliseconds. Mercury-wetted switches canalleviate this problem to a degree.

For high-voltage switching in air, a gap made of twopointed electrodes can be used as the discharge switch.Place the tips pointing toward each other and adjust thedistance between them to about 1 mm/kV of dischargevoltage. The gap must be tested and adjusted before thetest, and it must be verified that breakdown occurred at

the desired voltage. For tests that involve a variableamplitude, a safety gap connected in parallel is suggested.The second gap should be securely set at the maximumpermissible test voltage. The primary gap can be adjustedduring the test from zero to the maximum voltage desiredwithout fear of inadvertent overtesting. The test isperformed by charging the capacitor (or triggering thespark coil) and relying on the spark gap to discharge atthe proper voltage.

The arc source’s power supply must be sufficientlyisolated from the discharge so that the discharge is atransient and not a continuing arc discharge. Aconvenient test rate is once per second. To accomplishthis rate, it is convenient to choose the capacitor andisolation resistor’s resistance-capacitance time constantto be about l/2 second and to make the high-voltagepower supply output somewhat higher than the desireddischarge voltage.

For tests that involve a fixed discharge voltage, gasdischarge tubes are available with fixed breakdownvoltages. The advantage of the gas discharge tube overneedlepoints in air is its faster risetime and its veryrepeatable discharge voltage. The gas discharge tube’sdimensions (5 to 7 cm or longer) can cause more radio-frequency radiation than a smaller set of needlepoint airgaps*

Another type of gas discharge tube is the triggered gasdischarge tube. This tube can be triggered electronically,much as an SCR can be turned on by its gate. Thismethod has the added complexity of the trigger circuitry.Additionally, the trigger circuitry must be properlyisolated so that discharge currents are not diverted by thetrigger circuits.

4.3.2 Methods of ESD applicationThe ESD energy can range from very small to large (as

much as 1 J but usually millijoules). The methods ofapplication can range from indirect (radiated) to direct(applying the spark directly to a piece part). In general,the method of application should simulate the expectedESD source as much as possible. Several typical methodsare described here.

4.3.2. I Radiated field tests.-The sparking devicecan be operated in air at some distance from thecomponent. This technique can be used to check for rfinterference to communications or surveillance receiversas coupled into their antennas. It can also be used tocheck the susceptibility of scientific instruments that maybe measuring plasma or natural radio waves. Typical rf-radiated spectra are shown in figure 12.

4.3.2.2 Single-point discharge tests.-Discharging anarc onto a spacecraft surface (or a temporary protfdwmetallic fitting), with the arc current return wire in closeproximity, can represent the discharge and local flowingof arc currents. This test is more severe than the radiated

21

220 r

601 I I I I I.Ol .l 1 10 loo loo0Frequency, MHz

Figure 12.-Typical.1541 arc source.

radio frequency-radiated fields from MIL-STD-

test, since it is performed immediately adjacentrather than some distance away.

to the

This test simulates only local discharge currents; it doesnot simulate “blowoff” of charges, which causescurrents in the entire structure of the spacecraft.

4.3.2.3 Structure current tests,-The objective ofstructure current testing is to simulate “blowoff” ofcharges from a spacecraft surface. If a surface charges,and a resultant ESD occurs, the spark may vaporize andmechanically remove material and charges without localcharge equalization. In such a case the remaining chargeon the spacecraft will redistribute itself and causestructural currents.

Defining the actual blowoff currents and the paths theytake is difficult. Nevertheless, it is appropriate to do astructure current test, to determine the spacecraftsusceptibility, by using test currents and test locationssupported by analysis as illustrated in sections 2.2 and2.3. Typically, such a test would be accomplished byusing one or more of the following current paths (fig. 13):

(1) Diametricallyspacecraft)

locations (through the

Figure 13 .-Pathsstructure.

for electrostatic discharge currents through

(2) Protuberances (from landing foot to top fromantenna to body, and from thruster jets to opposite side.of body)

(3) Extensions or booms (from end of sensor boom tospacecraft chassis and from end of solar panel tospacecraft chassis)

(4) From launch attachment point to other side ofspacecraftThe test using current path 1 is of general nature. Testsusing current paths 2 and 3 simulate probable arclocations on at least one end of the current path. Thesetest points include thrusters, whose operation can triggeran incipient discharge, and landing feet and theattachment points, especially if used in a dockingmaneuver, when they could initiate a spark to the matingspacecraft.

Test 3 is an especially useful test. Solar panels oftenhave glass (nonconductive) coverslides, and sensors mayhave optics (nonconductive) that can cause an arcdischarge. In both cases, any blowoff charge would bereplaced by a current in the supporting boom structurethat could couple into cabling in the boom. Thisphenomenon is possibly the worst-case event that couldoccur on the spacecraft because the common length ofthe signal or power cable near the arc current is thelongest on the spacecraft.

4.4 Unit Testing4.4.1 General

Unit ESD testing serves the same purpose it serves instandard environmental testing (i.e., it identifies designdeficiencieg at a stage when the design is more easilychanged). However, it is very difficult to provide arealistic determination of the unit’s environment ascaused by an ESD on the spacecraft.

A unit testing program could specify a single ESD testfor all units or could provide several general categories oftest requirements. The following test categories areprovided as a guide:

(1) internal units (general) must survive, withoutdamage or disruption, the MIL-STD.1541 arc source test(discharges to the unit but no arc currents through theunit’s chassis).

(2) External units mounted outside the Faraday cage(usually exterior sensors) .must survive theMIL-STD-1541 arc source at a S-kV level with dischargecurrents passing from one corner to the diagonallyopposite corner (four pairs of locations).

(3) For units near a known ESD source (solar cellcoverslides, Kapton thermal blankets, etc.), the sparkvoltage and other parameters must be tailored* to besimilar to the expected spark from that dielectric surface.

22

4.4.2 Unit test configurntion

.

ESD tests of the unit (“subsystem”) can be performedwith the subsystem configured as it ‘would be for astandard electromagnetic compatibility (EMC) radiatedsusceptibility test. The unit is placed on and electricallybonded to a grounded copper-topped bench. The unit iscabled to its support equipment, which is in an adjacentroom. The unit and cabling should be of flight construc-tion with all shields, access ports, etc., in flight condition.All spare cables should be removed.

4.4.3 Unit test operating modesThe unit should be operated in all modes appropriate

to the ESD arcing situation. Additionally, the unit shouldbe placed in its most sensitive operating condition(amplifiers in highest gain state, receivers with a veryweak input signal) so that the likelihood of observinginterference from the spark is maximized. The unitshould also be exercised through its operating modes toassure that mode change commands are possible in thepresence of arcing.

4.5 Spacecraft TestingThe system level test will provide the most reliable

determination of the expected performance of a spacevehicle in the charging environment. Such a test shouldbe conducted on a representative spacecraft beforeexposing the flight spacecraft to insure that there will beno inadvertent overstressing of flight units.

A detailed test plan must be developed that defines testprocedures, instrumentation, test levels, and parametersto be investigated. Test techniques will probably involvecurrent flow in the spacecraft structure. Tests can beconducted in ambient environments, but screen roomswith electromagnetic dampers are recommended.MIL-STD.1541 system test requirements and radiatedelectromagnetic interference testing are considered to bea minimal sequence of tests.

The spacecraft should be isolated from ground.Instrumentation must be electrically screened from thedischarge test environment and must be carefully chosenso that instrument response is not confused withspacecraft response. The spacecraft and instrumentationshould be on battery power. Complete spacecrafttelemetry should be monitored. Voltage probes, currentprobes, E and H shield current monitors, and othersensors should be installed at critical locations. Sensordata should be transmitted with fiber optic data links forbest results. Oscilloscopes and other monitoring instru-ments should be capable of resolving the expected fastresponse to the discharges ( 5250 MHz).

The test levels should be determined from analysis ofdischarging behavior in the substorm environment. It isrecommended that full level testing, with test margins,be applied to structural, engineering, or qualificationmodels of spacecraft with only reduced levels applied toflight units. The test measurements (structural currents,harness transients, upsets, etc.) are the key systemresponses that are to be used to validate predictedbehavior.

4.5.1 GeneralSpacecraft testing is generally performed in the same

fashion as unit testing; a test plan of the following sort istypical:

(1) The MIL-STD-1541 radiated test is appliedaround the entire spacecraft.

(2) Spark currents from the MIL-STD-1541 arcsource are applied through spacecraft structure fromlaunch vehicle attachment points to diagonally oppositecomers.

(3) ESD currents are passed down the length of boomswith cabling routed along them (e.g., sensor booms orpower booms). Noise pickup into cabling and circuitdisruption are monitored.

(4) Special tests are devised for special situations. Forexample, dielectric regions such as quartz second-surfacemirrors, Kapton thermal blankets, and optical viewingwindows should have ESD tests applied on the basis oftheir predicted ESD characteristics.

4.5.2 Spacecraft test configurationThe spacecraft ESD testing configuration ideally

simulates a 100 percent flight-like condition. This may bedifficult because of the following considerations:

(1) Desire for ESD diagnostics in the spacecraft(2) Nonfunctioning power system(3) Local rules about grounding the spacecraft to

facility ground(4) Cost and schedules to completely assemble the

spacecraft for the test and later disassemble it(5) The possible large capacitance to ground of the

spacecraft in its test fixture(6) ESD coupling onto nonflight test cabling

4.5.2.1 Zest diagnostics.-To obtain more infor-mation about circuit response than can be obtained bytelemetry, it is common to measure induced voltages dueto the ESD test sparks at key circuits. If improperlyimplemented, the very wires that access the circuits andexit the spacecraft to test equipment (e.g., oscilloscopes)will act as antennas and show noise that never would bepresent without those wires.

23

Two approaches have been used with some success.The first is using conventional oscilloscope probes, withgreat care. Long oscilloscope probes (3 m) were procuredfrom Tektronix. For the circuits being monitored, a small“tee” breakout connector was fabricated and inserted at‘I he’connecter nearest the circuit. Two oscilloscope probeswere attached to each circuit’s active and return wires andthe probe tips were grounded to satellite structure in theimmediate vicinity of the breakout tee. The probegrounds were less than I5 cm from the probe tip. Thesignal was measured on a differential input of theoscilloscope. Before installation the probes werecapacitively compensated to their respective oscilloscopep&amplifiers, and it was verified that their common-’mode voltage rejection was adequate (in short, normal

’ good practice). The two probe leads were twisted togetherand routed along metal structure inside the satellite untilthey could be routed out of the main chassis enclosure.They were then routed (stil1 under thermal blankets) ’along structure to a location as remote as possible fromany ESD test location and finally were routed to theoscilloscope. The oscilloscopes were isolated frombuilding ground by isolation transformers. Clearly, thismethod permits monitoring only a few circuits.

A second method of monitoring ESD-induced voltagewaveforms on internal circuits is the use of battery-powered devices that convert voltages to light-emittingdiode (LED) signals. The LED signals can be transmittedby fiber optics to exterior receiving devices, where thevoltage waveform is reconstructed. As with theoscilloscope probes, the monitoring device must becarefully attached to the wires with minimal disturbanceto circuit wiring. The fiber optics cable must be routedout of the satellite with minimal disturbance. Thedeficiency of such a monitoring scheme is that thesending device must be battery powered, turned on, andinstalled in the spacecraft before spacecraft buildup andmust operate for the duration of the test. ‘The need forbatteries and the high power consumption of LED’sseverely restrict this method.

Another proposed way to obtain circuit responseinformation is to place peak-hold circuitry at key circuits,installed as described above. This method is not veryuseful because the only datum presented is that a certainpeak voltage occurred. There is no evidence that the ESDtest caused it, and there is no way to correlate that voltage

with any one of the test sequences.such information is worthless.

For analysis purposes,

4.5.2.2 Nonfunctioning power system.--Spacecraftusing solar cells or nuclear power supplies often must usesupport equipment power supplies for ground testactivities and thus are not totally isolated from ground.In such cases the best work-around is to use an isolatedand balanced output power supply with its wires routedto the spacecraft at a height above ground to avoid straycapacitance to ground. The power wires should beshielded to avoid picking up stray radiated ESD noise;the shields should be grounded at the support equipmentend of the cable only.

4.5.2.3 Facility grounding.-To simulate flight, thespacecraft should be isolated from ground. Normal testpractice dictates an excellent connection to facilityground. For the purposes of this test a temporary groundof 0.2 to 2 MQ or more will isolate the spacecraft for thepurposes of the ESD test. Generally 0.2 to 2 MQ issufficient “grounding” for special test circumstances oflimited duration and can be tolerated for the ESD test.

4.5.2.4 Cost and schedules to assemble anddisassemble spacecraft.-Often testing is done in themost compact form possible, attempting to interleaveseveral tasks at one time or to perform tasks in parallel.This practice is incompatible with the needs of ESD-testing and must be avoided. A thermal vacuum test, forexample, is configured like the ESD test, but hasnumerous (nonflight) thermocouple leads penetratingfrom the interior to the exterior of the spacecraft. Theseleads can act as antennas and bring ESD-caused noise‘into satellite circuitry, where it never would have been.

4.5.2.5 Spacecraft capacitance to ground duringtest.-If stray capacitance to facility ground is presentduring the ESD test, it will modify the flow of ESDcurrents. For an better test the spacecraft should bephysically isolated from facility ground. It can be shownthat raising a 1.5m-diameter spherical satellite 0.5 m offthe test flooring reduces the stray capacitance nearly tothat of an isolated satellite in free space. A dielectrici(e.g., wood) support structure can be fabricated for theESD test and will provide the necessary capacitiveisolation.

4.5.2.6 ESD coupling onto nonflight testcabling.-One method of reducing ESD coupling to and

24

from the spacecraft on nonflight test wiring is the use offerrite beads on all such wiring.

.

An anomaly is known only to have occurred at sometime. Since most spacecraft are not well instrumented forenvironmental effects, the state of the environment at thetime of the anomaly would have to be inferred from

5.0 Control and Monitoring Techniques5.1 Active Spacecraft Charge Control

Charge control devices are a means of controllingspacecraft potential. Various active charged-particleemitters .have been and are being developed and showpromise of controlling spacecraft -potential in the spaceplasma. environment. At this time only neutral plasmadevices (both ion and electron emitters) havedemonstrated the ability to control spacecraft potential ingeomagnetic substorms. These devices are recommendedfor charge control purposes (Purvis and Bartlett, 1980;Olsen, 1978; Olsen and Whipple, 1977).

Emitted particles constitute an additional term in thecurrent balance of a spacecraft. Because the ambientcurrent densities at geosynchronous altitude are quitesmall, emitting small currents from a spacecraft can havea strong effect on its potential, as has been demonstratedon the Applications Technology Satellites ATS-5 andATS-6, Spacecraft Charging at High Altitudes(SCATHA), and other spacecraft. However, devices thatemit particles of only one electric charge (e.g., electrons)are not suitable for active potential control applicationsunless all spacecraft surfaces are conducting. Activationof such a device will result in a rapid change of spacecraftpotential. However, differential charging of anyinsulating surfaces will occur and cause potential barrierformation near the emitter. Emission of low-energyparticles can then be suppressed. Higher energy particlescan escape, but their emission could result in the buildupof large differential potentials. On the other hand,devices that emit neutral plasmas or neutralized beams(e.g., hollow cathode plasma sources or ion engines) canmaintain spacecraft potentials near plasma ground andsuppress differential charging. These are therefore therecommended types of charge control devices.

ground observatory data. These environmental data arenot necessarily the same at the. spacecraft location; infact, the correlation is generally poor.

This unknown condition could be modified if space-craft carried a set of spacecraft charging effect monitors.A simple monitor set has been designed that will measurethe characteristic energy and current flux of theenvironment as well as determine transients on fourharness positions within the spacecraft (Sturman, 1981).This will alIow correlation between the onset of thesubstorm environment and possible transients induced onthe electronic systems. This package weighs about 1.4 kgand uses less than 3 W of power. The environmentsensors would have to be on the outside surfaces andpreferably in shade.

More sophisticated packages are available. Ion particledetectors in the range 10 to 50 keV could be used to sensethe onset of geomagnetic substorms. Transient monitorscapable of measuring the pulse characteristics are alsoavailable (Koons, 1981). These would require largerweight and power budgets but do provide better data.

These spacecraft charging effect monitors require dataanalysis support in order to produce the desired results. Ifthey were carried on a number of operational satellites,the technology community would be able to obtain astatistical base relating charging to induced transients.The operational people, on the other hand, would be ableto tell when charging is of concern, to establish proce-dures minimizing detrimental effects, and to separatesystem malfunctions from environmentally - inducedeffects.

It is recommended that monitor packages be carried onall geosynchronous spacecraft. These packages must con-sist of both environment and transient pulse detectors.

5.2 Environmental and Event MonitorsThe occurrence of environmentally induced discharge

effects in spacecraft systems is usually difficult to verify.

National Aeronautics and Space AdministrationLewis Research Center .Cleveland, Ohio, April 16, 1984

Appendix A.Description of Geosynchronous Plasma Environments

. .

.

‘In section 2.0 of this document geosynchronousplasma environments are briefly introduced and simplydescribed in terms of temperature and number density.The environmerit is actually very complex and dynamicand not yet fully understood even by researchers, Thesimple characterization of the environment in section 2.0uses only two species, electrons and protons; it assumes a“single Maxwellian” distribution, where the energydistribution of each species is considered to be describedby the mathematical function, the “Maxwellian.” TheMaxwellian treatment is used because the function can beeasily treated in the necessary mathematicalmanipulations for- calculating spacecraft charging. If aMaxwellian is not used, measured data must be curve fit

digitally and at much greater computational cost. If asingle-Maxwellian distribution is inadequate for a givencircumstance, the measured data ark often treated as thesum of two populations, each with . a Maxwelliandistribution: the “two Maxwellian” characterization.Other species such as oxygen and helium can be treated asadditional Maxwellian populations. The following textdescribes in greater detail different characterizations ofthe geosynchronous plasma environment.

Characterizations of Geosynchronous PlasmaEnvironment

An initial step in looking at an environment is to *consider averages. Ten-minute averages of approxima@y45 days (per spacecraft) were made of data for theATS-5, - ATS-6, and SCATHA (experiment SC9)spacecraft (table VII); isotropy wals assumed. Thestandard deviations present in the data were estimated(table VIII). The ions were assumed to be protons in thesetables. Note that in many cases the standard deviationexceeded the average. This resulted from the .greatvariability of the geosynchronous environment andillustrates the inherent difficulty of attempting tocharacterize the “average” plasma environment. Thesevalues are useful, however, in estimating the prebtormconditions that a spacecraft will experience. As the initialcharge state of a spacecraft is important in determininghow the vehicle will respond to a significant environ-mental perturbation, this is useful information. Alsothese averages give an approximate idea of how plasmaconditions vary over a solar cycle since the ATS-S dataare for 1969-70, the ATS-6 data for 1974-76, and theSCATHA data for 1978.

. .

TABLE VII. - AVERAGE PARAMETERS

Parameter

ATS-5 ATS-6 SCATHA

Number density, (ND), cmm3Current density, (J), nA cnr2Energy density, (ED), eV gm3.Energy flux, (EF), eV cm- so1 sr-lNumber density for population 1, N , cr3Temperature for population 1, Tl, rt eVNumber density for population 2, NTemperature for population 2, T2, z

, cmo3eV

Spacecraft

0.800.068

19E,98x100.5780.2770.2157.04

1.060.09635

2.17x10990.7510.4600.2739.67

1.090.115

37E!1.99x100.7800.5500.3108.68

Average temperature, Tav, keVRoot-mean-square temperature, T,,, keV

(b) Ions

2

1.305.1

l3 OPP'~6x100.750.300.6114.0

Average temperature, Tav, keVI

6.8Root-mean-square temperature, T,$, keV 12.0

1.203.4

l2 OpP3.4x100.930.270.3325.0

12.023.0

0.583.3

9442.0x10 1P

0.190.800.3915.8

11.214.5 I

26

TABLE VIII. - STANDARD DEVIATIONS

(a) Electrons

Parameter Spacecraft

ATS-5 ATS-6

Number density, (ND), cm=3Current density, (J), nA CIK~Energy density, (ED), eVEnergy flux, (EF), eV cm- 5

m3s-1 srl

* -+31+1,7Yl0 !P

Number density for population 1, NTemperature for population 1, Tl, t LV

cm-3 - +0.55To.17

Number density for population 2, Nz

, cm-3 q.38Temperature for population 2, T2, eV -+2.1

Average temperature, Tav, keV +2.0Root-mean-square temperature Trms, keV T3.3

L

+l.l+n.os737

+2.6x0 !!i!+0.82TO.85To.3413.3.6-

‘2 . 0T3.5

(b) Ions

Number density, (ND), cm-3Current density, (J), pA cm-2

+0.69

Energy density, (ED), eV Em-3-+2.7

Energy flux, (EF), eV cm- s-1 w-1+v7

+3.55110 PPNumber density for population 1, N

It, cmo3 - +o.M

Temperature for population 1, Tl, eV To.30Number density for population 2, N

z, cm-3 x33

Temperature for population 2, T2, eV -+5.0

Average temperature, Tav, keV +3.6Root-mean-square temperature, TrmS, keV T4.8

+1.7T1.8

+v1+3.6ylOPp

+1.78TO.88TO.16-+8.5

+8.4- I

+4.6T8.9 T5.3-

A second way of considering environments is to look at“worst case” .situations. Worst-case estimates of theparameters in table VII were made for the geosyn-chronous environment (table IX). These values werederived from fits to the plasma distributions observedduring the several known worst-case charging events. TheSCATHA spacecraft instrumentation allowed a breakoutof the data into components parallel and perpendicular tothe magnetic field and thus permited a more realisticrepresentation of the actual environment. These valuesare particularly useful in estimating the extremes inenvironment that a geosynchronous spacecraft is likely toencounter.

A third quantity of interest in estimating the effects ofthe space environment on charging is the yearlypercentage of occurrence of the plasma parameters. Theoccurrence frequencies of the temperature and current(fig. 14) were derived by fitting the observed distributionsof electron and ion temperature for University ofCalifornia at San Diego instruments on ATS-5, ATS-6,and SCATHA. The current values were computed fromthese latter curves by assuming an adiabatic relationship.The figures should be useful in estimating the time duringthe year that a specified environment might be expected.

The fourth and a very important quantity of interest ishow the plasma parameters vary with time during acharging event. The approaches to determining this quan-tity range from detailed models of the magnetosphere toaverages over many geomagnetic storms. For design

* - 1SCATHA 1

+0.89TO.10734

+2.oaO98I- +0.70TO.32a37-+4;0

+1.5rZ.9

purposes we have adopted a simulation of the electronand proton current and temperature that approximatesnatural variations in the potential as predicted bycharging analysis codes (e.g., NASCAP). A timevariation sequence suitable for modeling the worst effectsof a geomagnetic storm is recommended (fig. IS). (Notethat the simple single-Maxwellian representation has beenfound to match flight data when used in NASCAPstudies.)

Derivation of Moments of Plasma Distribution Function

The Earth’s plasma can be described, as discussedearlier, in terms of simple Maxwell-Boltzmanndistributions. As this representation lends. itself toefficient manipulation when carrying out chargingcalculations, it is often the preferred way for describingplasmas. The Maxwell-Boltzmann distribution Fi is givenbY

F,(v) = ni ( &.)3’2exP(s)

where

ni number density of species imi mass of species ik Boltzmann constantTi temperature of species i

(Al)

27

TABLE IX. - WORST-CASE GEOSYNCHRONOUS ENVIRONMENTS

[The moments Tav and Trms are averaged over al 1 angles. .and perpendicular to the magnetic field.

The SCATHA two-Maxwellian parameters are for fluxes parallelATS-6 two-Maxwellian parameters are averaged over all directions.]-

Parameter SourceL

Deutsch (1981) Mullen et al. Mu1 len and(1981) Gussenhoven (1982)

Date,

\ Day 178, 1974. Day 114, 1979 --a

Spacecraft *

ATS-6 SCATHA SCATHA

Electrons Ions Electrons . Ions Electrons Ions

Number density, (ND , cdnA cnr2

1.12 2.3(3

0.245 0.900 !I 3.00 3.00Current density, ,Energy denisty, (ED), eV $r3

0.41!0.293x1

0.025f 0.184 0,795x10- 0.505 0.015~Energy flux, (EF), eV CC s-l srl 0.264x10 4P 0.104x1

0.298x10 P2 o.960x1p3 0.370x10.668x10

0.19Oxlp:0.430x10

0.240xlf40.151x10 0.748~10 P2

Number density for population 1, Nl, cm-3:Parallel 0.882x10-2 . 0.200 1.60 1.00 1.10Perpendicular - - - - 0.200 1.10 0.800 0.900

Temperature for population 1, Tl, eV:Parallel :

0.111x103 0.400x103 0.300x103 0.600x103 0.400x103Perpendicular O - - W & -

Number density for population 2, N2, cmo3:0.400x103 0.300x103 0.600~10~ 0.300x103

P aral 1 el 1.22 0.236 0.600 0.600 1.40 1.70Perpendicular - - - 2.30 1.30 1.90 1.60

Temperature for population 2, T2, eV:Parallel 0.160x105 0.295~10~ 0.240~10~

0.248~10~0 .260x1050.282~10~

0.251x1050.261~10~

0.247x105Perpendicular - - - - - - - 0.256~10~

Average temperature, Tav, ‘eV ’ 0;160x1050.161~10~

0.284~10~Root-mean-square temperature, Tms, eV . 0.295x105

0.77oxq4 0.550x1040.140x105

0.533~1~0~ . 0.822~1040.900x10 . 0.733x104 0.118~10~

V velocityFi distribution function of species i S

Unfortunately, the “Bpace plasma environment isseldom a Maxwell-Boltzmann distribution. However,given the actual plasma distribution function, it ispossible to define (irrespective of whether the plasma isMaxwell-Boltzmann or not) moments of the distributionfunction that reveal characteristics of its shape. Thesemoments can in most cases then be used to determine anapproximate Maxwell-Boltzmann distribution. The frstfour of these moments are

whereI ..

WDi) first moment that equals ni(NFi) number flux of species i .-.(EDi) energy density of species iW)i energy flux of species i . .For the Maxwell-Boltzmann distribution of equation(Al) these assume the following values:

(NDJ = 4~ srn (Uo) Fiv2dv0

(NFJ = j- (vl)Fiv2dv0

Often it is easier to measure the four moments of the

WVi=-mi I=+ (v3)F#dv2 VW0

plasma distribution function than the actual temperature.This & particularly true for space plasnias, where theconcept of “temperature” is not well defmed. As an

28

. 2.. : .ii)r

. z%o,-v

lidld-

k Electron current, nA/cmz

C

-

ld ld lob . ldProton current r&m* ’ Pruton tempature, -eV

Figure 14.-Occurrcn~ frequencies of plasma parameters.

16 -

14 -

=>,I*- .YfiJ .g lo.-85 8-CI

.. 6 -w-

4 -

2 -

Time, min

2.8

2.4

. . F~urq 15.‘Tie history of model substorm..

29

illustration, from the four moments, two definitions of”the plasma temperature can be developed:

T WD)a v= T(NQ

l *. .

*.

.

W tTMIS =-

(NV

. ;(ilO)

(Al 1)

For a true Maxwell-Boltzma& plksma there qua&&swould be equal; for act&L plasm& 7& is uklly greaterthan Tav* Even so, experience Aas ; shown that arepresentation in terms of two kxwell-Boltzmanndistributions is in fact a ‘better mathematicalrepresentation of the space plasma than a singleMaxwellian. That is, the plasma distribution for a singlespecies can be represented by :’ .

; .

where .< *

N1 number density for population 1T1 temperature for population 1N2 number density for population 2T2 temperature for population 2

This representation in most cases fits the data quiteadequately over the energy range of importance tospacecraft charging. Further, it is very simple to derive. .N,,’ ?I; -N2, and .Q directly .from the four moments sothat a consistent mathematical representation of theplasma [can be estabhshed that incorporates the simplicityof the Maxwell-Boltzmann representation whilemaintaining a physic&y reasonable picture of theplasma. The distinction between Tav, Tms, Tl, and T2must be kept. ii mind, hoGever, whenever reference ismade to a “MaxwellBoltzmann” distribution as theseare only approximations at best to the actual plasmaenvironment.

. s-1: . . ..

* .-.

30

Appendix BTechnical Description of NASCAP. *

The NASA Charging Analyzer Program (NASCAP) isa quasi-static computational program (i.e., it assumesthat currents are functions of environmental parameters,electrostatic potentials, and magnetostatic fields but arenot dependent on electrodynamk effects). This isreasonable since charging times in insulators are- longcompared with the computing interval. The followingparagraphs briefly discuss the elements of NA$CAP.Detailed descriptions (Katz et. al., 1977, 1979, 1981),including a users manual (Cassidy, 1978), are available.

A flow diagrani of NASCAP is shown in figure 16. Thelogic h& been designed to provide maximum flexibility tothe user.’ As execution progresses, the user may request acharging simulation or any of several auxiliary functions

I Objectdefinition I

Initialpotentials

II I

-I Trajectorycalculations I

77 (Incoming flux)

Surface’ ’interactions -

(Charge distribution 1

‘7; (Potentlall

Figure 16.-Flow diagram of NASA Charging Analyzer Program(NASCAP).

such as object definition, particle emitter, or detectorsimulation. NASCAP contains full restart capability.

A NASCAP charging simulation first calculates thecurrents incident on and emitted from all surfaces for agiven environment. From these currents it computes thenew electrostatic potenti& (relative to the space plasmapotential) on all spacecraft surfaces and in surroundingspace. This iteration continues for a user-specified periodof time. The charging simulation can take into accountsuch effects as internal bias voltages, Debye screening,and charged-particle emitters.

Computational SpaceNASCAP computations are performed in an embed-

ded set of cubic grids of dimensions 17 x 17 x (4n + l),where 4z~n ~8 (fig. 17). The object is described in theinnermost grid. Each successive grid has twice the lineardimensions of the next inner one. This alloys treatmentof a large volume of space while minimizing compu-tational time and storage.

Environment DefinitionIn NASCAP the charged-particle environment can be

specified in a number of ways. For simulating’ spaceenvironments the most commonly used techniques are theMaxwellian and double-Maxwellian descriptions ofgeomagnetic substorms. These allow independent speci-fication of temperatures and particle densities of bothelectron and proton components. Actual particledistributions from flight instruments can also be used tospecify the space environment for both quiescent andsubstorm conditions. Anisotropic fluxes to varioussurfaces can also be defined.

Third grid

Figure 17.-NASCAP nested grid computational space.

NASCAP also treats surface charging in laboratory’simulations. For these cases, single or multiple beams of .electrons or ions can be specified from’ arbit&y”’locations. Since the surface charging physics is the samefor both space and laboratory simulation environments,the accuracy of NASCAP predictions can be determined .from laboratory results’where environment fluxes can becontrolled ;ind detaued measurements made. Thepredictions are“ generally within ’ 20 percent of theexperimental surface’voltages.

Object Definition ’ . -, .NASCAP require;.that an object be defined in terms of W

thin booms, flat plates, rectangular parallelepipeds, orsections of parallelepipeds. Only thin booms extendbeyond the innermost grid boundary. Furthermore, noother portion of the object must touch.the innermost grid ”boundary. This object definition protocol allows rather.complex spacecraft models to be defined by using fairly .simple inputs j.

-. t . . ‘.. .

Since a spacecraft can be a complex shape and errors indescribing the model in terms of program limitations canarise, a graphical output of the spacecraft model can begenerated by the computer to verify the accuracy of themodel before start of computatiofis. (See fig. I& for anexample of output.) Any set of axes or rotational angles . .can be specified for viewing the object. The graphicaloutput of the object. definition identifies the specifiedmaterials used on the surfaces. Hence, it is possible todetermine that the computer model is the desired.representation of.the spacecraft. ‘ .*.

Material PropertiesNA$CAP allows surfaces to be bare or covered with a

thin (- 10-4 m) dielectric material. Values of propertiesfor common spacecraft materials (e.g., aluminuin, gold,Teflon, Kapton, and silica) are supplied in the code and

can be adjusted if desired by the user. Properties forother materials must be specified by the user. Theproperties required ’ by NASCAP are dielectric constant,material thickness, backscatter and secondary emissioncoefficients (for .both electron and proton impact), bulk. .a n d surf&e cond’uctivit~es,’ photoe&ssion yield,.ele&ical breakdown thresholds, and r&iatid&ndu~ed .. .conductivity propert ies: . ’ *

. . ... . . .’ . ‘.i. . . . .* -- 1

Electrical Conaectivity . ” . .. . . * . -I . 1.. .

In NASCAP the’ spacecraft model‘cari be &posed of’. -up ‘to 15 sepaiate cdnductors. These conductors & beresistively or capacitively coupled and can be allowed tofloat, to be held ‘at fixed ‘potentials,’ or ‘to’ be biased-.relative to one &&her. In the latter kase, HASCAPautoniatically transports charge from one conductor to

.:

another to main& the bias voltiges:- . . .* -

Mathematical AlgorithmNASCAP uses an incomplete Cholesky conjugate .

gradient algorithm to calculate the change in spacecraftpotential at each time step (- 103 variables). Thespacecraft equivalent circuit used in this calculation is setup by geometrical analysis within NASCAP. Thepotential in the external space (- 104 to 10s variables) iscalculated by a finite-element, sealed-conjugate-gradienttechnique. Both potential solvers are capable of handlingmixtures of fixed-potential and fix&charge boundaryconditions at the spacecraft surface.

DetectorsAt any time interval after the charging simulation

begins, the user can request a simulation of a particledetector behavior. The user specifies the location of adetector, an aperture, and a range of viewing angles orparticle angles. NASCAP then computes particle

Mi(si

Cutaway view showingoptics system

. , .* :CS-809594 .

Figure 18.- NASCAP model of large optics system satellite.

32

trajectories, by using predicted surface voltages andexternal fields, from the detector location to eitheremission from another part of the spacecraft or arrivalfrom space. (See fig. 19.) Those particles arriving fromspace are assumed to be those that the particle detectorshould sense. Both electron and proton detectors can bespecified. . *. . . . . .

outputIn addition to its standard printed output, NASCAP

provides an extensive menu. of graphical outputs and

Figure lg.-Electron trajectories for Galileo (Hare1 et al., 1982).

printed data compilations. Graphical output includes thematerial and perspective object definition pictures,potential contour plots, and particle trajectory plots. Thestandard printed output includes a summary of all cellvoltages, listing of currents to specified surface cells, andcompilation of electrical stress through insulators indecreasing order. Sorting routines can tabulate specificcell potentials as a function of time for specified sets ofcell numbers or materials. Sufficient information isstored in external files to allow a restart of a NASCAPprogram for further analysis, for evaluation underchanged environmental conditions, or for postprocessinganalysis with user-written programs.

33

.

. Appendix CVoyager SEMCAP Analysis . .

To simulate the effects of arc discharges on Voyager,either a high-voltage-excited spark gap or a flat-platecapacitor with an arc gap was used to induce arcs. Theradiated fields from these sources were approximated inthe Specification and p Electomagnetic CompatibilityProgram (SEMCAP) in two ways:

(1) An induction field model consisting of quasi-staticelectric and magnetic. fields proportional to the voltageand current of the source, respectively ’

(2) A radiated field model representing the far-fieldelectromagnetic radiation of the loop antenna formed bythe source ’

The Voyager test results using SEMCAP and theseassumed arc source ‘models are presented in table Xalong with the values actually observed. The source

. . . .* . 4 .

*. . . ) I . .’ , l . ;

\

.. . . .parameters used in the predictions are’ presented in tableXI (reflecting the arc parameters of the test source). Themean error between the predicted and measured results is-6 dB, and the standard deviation is 23 dB. Assuming. . .these accuracy parameters to be applicable to predictedin-flight responses for Voyager, the space& &asconsidered to be immune to arc discharges below 20 mVon the basis of the SEMCAP analysis. The use ofSEMCAP in this application caused numerous designchanges that significantly improved the arc dischargeprotection of the Voyager spacecraft. Even though flightVoyagers still suffered several arc discharge events, thedesign changes resulting from SEMCAP (in conjunctionwith testing) are believed to have significantly enhancedtheir survivabjlity. ., .

.. . I.* . . .

. 6.

t-. l:: .

. .

: ; TABLE X. - SEMCAP PREDICTIONS.* -

[Mean error, -12 dB (underpredicting); standard deviation, 20 dB - not including entries footnoted a.].Location of arc .

. .High gain antenna Infrared interferometer/ OBLFM Sun sensor

spectrometer

v . 7. \ . . Predicted Measured Predicted Measured Predicted. Measured Predicted Measured.. ... a . I. Voltage values from radiated tests

Infrared interferometer/spectrometer 0.9 1.2 2.0 0.8 0.09 ' 1.0 0.08(IRIS) imaging subsystem

2.5

Low-energy charged-particle experiment 1.5 .8 6.0 a .45 .4 .7 .l 1 . 6Sun sensor . ' .84 1.0 .48 a .3 .9 a .4 4.0 1.6MagnetometerFrequency-selective subreflector

.4 a .4 .8 a .5 3.6 1.2 .4 .5

Brewster plateb4.9 4.0 .96 1.5 2.1 1.4 .002 2.02.2 4.0 1.5 4.0. .9 3.3 .54 f 17.7

I Voltage values from surface tests 1

Low-energy charged-particle experiment 0.27 0.6 15 a0.6 0.04 0.6 0.017 0.8Brewster plate 6.8 1.0 .37 .6 A .7 .4 .9Frequency-selective subreflectcr 57 Cl0 .09 c4.2 2 a1.7 .OOl a4.01 1 \aBackground noise; noise due to arc unnoticeable.bPredicted was *contact"-test; measured was "radiated" test.CExtrapolated.

TABLE XI. - IN-FLIGHT ARC MODELS - SEMCAP PREDICTIONS VERSUSESD TEST RESULTS (FLIGHT SPACECRAFT)

[Parameters of in-flight arc models (after scaling test data to spacecraft dimensions).]

. Arc source

Magnetometer cableHigh-gain antennapaint (outboard)

Plume shield .(sep-aration connector)

Frequency-selectivesubreflector

High-gain antennapaint (inboard)Plume shield (radio-i sotooe thermo-electric generator,RTG)

RTG bxideModified infraredinterferometer/spect?ometer (MIRISKaptonBrewster plateaSeparation connectorMagnetometer Teflonb

reakdownvoltage,

V,kV

ischargeor arccurrent,

1,*. A

ischargecurrentisetime,

tr9ns

5:l .

1 . . '

7

1

1

.’ 20 10150 * 5

16 20

80 8

150 5

16 20

3.51

925150

1 25 361 3

aNo area scaling needed because samplebTeflon models are believed to be well

was entire item.understood.

Dischargecurrent

ulse width,

ins

17003000

285

80

2400

330

Maindischargeapacitance,

C,.nF .

50 '400

4.5

,014

300 .

5.2

3700 34026 .04

10 2015 .1513 .038

Energy 9Lm3

62.5200

2.25

.34

.15

2.6

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40

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42

1 :Report No.

NASA TP-23614. Title and Subtitis

2. Government Accession No.

I

3. Recipient’s Catalog No.

5. Report Date

Design Guidelines for Assessing and ControllingSpacecraft Charging Effects

7. Author(s)

506-55-728. Performing Organization Report No.

Carolyn K. Purvis, Henry B. Garrett, A. C. Whittlesey,and N. John Stevens

9. Performing Organization Name and Address

National Aeronautics and Space AdministrationLewis Research CenterCleveland, Ohio 44135

m--p13. Type of Report and Period Covered

2. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D.C. 20546

5. Supplementary Notes

C. K. Purvis, Lewis Research Center; H. B. Garrett and A. Whittlesey, CaliforniaInstitute of Technology, Jet Propulsion Laboratory, Pasadena, California;N. John Stevens, Hughes Aircraft Company, El Segundo, California.

6. Abstract .

Experience has indicated a need for uniform criteria, or guidelines, to be usedin all phases of spacecraft design. Accordingly, guidelines have been developedfor the control of absolute and differential charging of spacecraft surfaces by thelower energy (less than approximately 50 keV) space charged-particle environment.Interior charging due to higher energy particles was not considered. Thisdocument is to be regarded as a guide to good design practices for assessingand controlling charging effects. It is not a NASA or Air Force mandatoryrequirement unless specifically included in project specifications. It isexpected, however, that this document, revised as experience may indicate, willprovide uniform design practices for all space vehicles.

17. Key Words (Suggested by Author(s))

Spacecraft chargingGeosynchronous spacecraft designElectromagnetic compatibility

- - -18. Distribution Statementr Unclassified. - Unlimited

STAR Category 18

19. Security Classif. (of this report) 20. Security Classif. (of this page)

Unclassified Unclassified

.21. No. of pages

4422. Price’

A03

l For sale by the National Technical Information Service, Springfield, Virginia 22161 NASA-Lang1 ey , 1984