Table of Contents ECSS Model Page
Background Information Background Information
Spacecraft charging

Table of contents

Introduction
The low Earth orbit environment
Spacecraft/plasma interactions
Evaluation of in-orbit spacecraft/plasma interactions
Introduction
Spacecraft wake effects
Spacecraft contamination effects
Spacecraft power system effects
Spacecraft electromagnetic effects
Evaluation of in-orbit spacecraft charging
Spacecraft charging effects
Conditions required for spacecraft charging
Summary of previous work
The ESPIRE computer codes
Scope
Description of the ESPIRE codes
LEOPOLD
SOLARC
EQUIPOT
SAPPHIRE
PICCHARGE
The DICTAT computer code
References

Introduction

The problems caused by interactions between a spacecraft and the environment in low Earth orbits and polar orbits (see Fig. 1 for a schematic representation of the polar orbit environment) are manifold ( , 199). Five broad areas of spacecraft/plasma interactions cause effects of concern for future space missions:

Many of these phenomena will be affected by the move to large, active structures, which will be accompanied in their orbits by a self-generated cloud of contamination. The trends to larger size, longer lifetimes, higher power requirement and frequent thruster firings and water and other waste dumps will pose a series of important problems for the space system designer and operator.

Schematic of interactions in
    polar orbit
Figure 1. Plasma interactions with a spacecraft in polar orbit

The low Earth orbit environment

The plasma environment in low Earth orbit (LEO) usually is relatively benign compared to the environment commonly found in geostationary Earth orbit (GEO). There is a large reservoir of high density cold plasma which tends to mitigate any spacecraft charging effects by providing a large source of charged particles from which neutralising currents may be drawn. The spacecraft is closely coupled to the plasma environment under most conditions.

However, since Debye lengths (plasma charge shielding distances) in the ionosphere are of the order of centimetres, as compared to metres or tens of metres at geostationary altitudes, the effective current collecting area of a spacecraft in LEO will be much smaller than that in GEO.

In addition, the spacecraft will create a large volume of disturbed plasma as it moves through the ionosphere. The spacecraft orbital velocity is hypersonic with respect to the plasma ions, and so phenomena such as bow shocks, ion voids, trailing wakes and density enhancements and depletions will occur. These aspects of the spacecraft/plasma interaction in LEO will not only cause a complicated, disturbed environment local to the spacecraft, but significant wake effects could introduce problems with differential charging.

This is liable to be a problem of some importance in the auroral precipitation zones at high magnetic latitudes. Here, energetic electrons, with energies typically in the range of 5-10 keV, precipitate down magnetic field lines in sufficient intensity to have the potential of greatly modifying the spacecraft charging situation.

Thus, in auroral regions a spacecraft moving along its orbit could suddenly experience a transition from a situation with a spacecraft charging voltage no more than a volt or two negative, to the case where it is intercepting a high energy stream of electrons. If this occurs in darkness, where photoemission is suppressed, and the electrons impinge on wake surfaces where ions are excluded by the spacecraft motion, then the potential for charging is high.

Spacecraft/plasma interactions

The main spacecraft/environment interactions are summarised in Table I, together with the effects they produce, and comments relating to their impact on large spacecraft systems (typical of the elements being developed as part of the International Space Station programme).

Table I. Summary of spacecraft/plasma interactions and their effects
InteractionEffectsImpact on large spacecraft systems
Supersonic motion through plasmaDensity enhancement at ram; density depletion in wake.Disturbances will be large in spatial extent; many body dimensions.
Collection of energetic auroral electronsNegative charging in wake; differential charging at surface.Absolute voltage levels will be much higher than for small or medium size bodies.
Dual spacecraft manoeuvresCharging of spacecraft in wake, if subject to energetic electron fluxAbsolute voltage levels on second body will be very high.
Neutral (natural and artificial atmosphereContamination of spacecraft surfaces; contamination of space environment.Greater levels of contaminant emission, and increased possibility of disturbance.
Ion attraction by surfaces at negative potentialSputtering of surface materialsLife-limiting factor for surfaces and solar array interconnects
Magnetic fieldModification of charging levels by suppressions of electron escapeAbsolute voltage levels will be much higher than for small or medium size bodies.
Motion in magnetic fieldGeneration of electric field gradientsAbsolute field values will be greater than for small or medium size bodies.
Current collection by solar arraysPower loss drain through plasma; arcing damage to surfaces.Power losses may become unacceptable.
Aspects of motion, spacecraft power leakage, contaminationGeneration of a range of plasma waves, at a range of frequenciesA wider range of generation mechanisms, with more power, and increased possibility of disturbance
Emitted particle beams in a plasmaBeam-plasma discharges, excitation of plasma waves, spacecraft charging, optical emissionsIncreased possibility of disturbance and enhancement of effects
Coupling of electromagnetic wavesLocalised heating of plasma at critical frequenciesIncreased possibility of disturbance and enhancement of effects

There are five broad areas where spacecraft/plasma interactions cause effects which may be of great concern for future missions.

  1. Effects caused by the supersonic spacecraft motion through the background ions in the plasma. Large density and space potential disturbances will be set up in the wake of bodies, and density enhancement will occur in the ram. These disturbances will be many body dimensions in spatial extent, and could lead to the shadowing of surfaces, disruption to instrumentation, enhanced particle fluxes to other parts of the spacecraft, and many other related effects. The motion across magnetic field lines will also introduce electric fields, and the effects of these will be greater for large structures.
  2. Solar arrays operating in the relatively dense plasmas in low Earth orbit will experience a current drain on the power system, as a result of losses through coupling to the plasma. This will become more severe as systems are enlarged and operating voltages are raised. Arcing of solar arrays in plasmas could lead to damage and electrical noise. Arcing becomes likely when the array is in a contaminated envrionment. Attraction of the ambient ions by negatively charged areas of the array could lead to sputtering and mass loss rates that may limit the lifetime of components, especially those designed for long duration operation.
  3. Contamination will be a critical issue in future missions, and this can be divided into two areas of concern. The first is that of spacecraft contamination, where material properties may be changed, thermal contral systems affected, delicate sensing equipment damaged, etc. The relatively dense atmospheric pressure in LEO is important here. The second aspect is that of modification of the ambient atmosphere by outgassing from the spacecraft structure, thruster firings, water dumps, etc.
  4. There will be large scope for the generation and emission of plasma waves, not just by active beam emissions experiments, but also by other spacecraft interactions. Plasma wakes, spacecraft power leakage and contaminant ions all provide a source of radiofrequency emissons. As spacecraft become larger and produce more contamination, the possibilities of disturbance, and of a wider range of generation mechanisms, frequencies and power levels, will all increase.
  5. Exposure to high energy auroral electron fluxes in the polar regions can lead to high levels of charging on spacecraft, particularl if the current collectoin occurs in the ion-depleted wake zones. The orientation of the magnetic field will complicate the interaction mechanisms, with fields parallel to the collection surface inhibiting the escape of secondary emitted electrons and enhancing charging. Again, these effects will be more important for large structures, where voltages well in excess of the kilovolt level have been predicted to occur. A related problem is that of the charging of one body immersed in the wake of another. The dynamic nature of the charging mechanisms will materials dependent and differential charging between adjacent surfaces could be of concern, especially for large structures.
It must be appreciated, however, that the interactions and effects referred to above are not to be looked at in isolation from one another. Many of the interactions not only produce primary effects, but are also responsible for secondary effects, acting in synergism with other interactons. As an example, a large structure with a large ion-free wake zone in a flux of energetic electrons could experience high negative charging levels. These in turn could attract contaminant ions to sensitive surfaces and modify the local plasma environment, causing plasma wave emissions which could disrupt sensors and monitors, or even in extreme cases on-board computers and communications.

Evaluation of in-orbit spacecraft/plasma interactions

Introduction

A spacecraft moving through the space plasma at LEO orbital speeds of the order of 8 km/s produces changes in the lcoal environment, and the local environment induces changes in the properties and performance of the spacecraft. These interactions result in significant changes in the local environment properties and include enhancement of neutral and ionised particle densities in the ram direction and rarefaction in the wake behind the spacecraft.

While ions and electrons will be constrained by the magnetic field of the Earth, neutral particles generated by the spacecraft will be free to travel with the vehicle until disturbed by collisional processes. This will be especially of concern for large structures carrying out active emissions of neutral particles, e.g. water and waste dumps, thruster firings, atmospheric venting, etc. In one sense, a spacecraft carries its own atmosphere along with it in orbit.

Space power systems will have to operate in a plasma environment. For solar arrays, with exposed solar cell interconnects, the ionised plasma may be a cause of current leakage and power loss from the array. Operation of the array at high voltage, to reduce cabling losses and array mass, may experience problems with plasma induced breakdowns and arcing, and with anomalous current collection phenomena. Biasing of interconnects at negative voltages in excess of sputter thresholds could lead to ion-induced erosion of the interconnects and destruction of the integrity of the array.

The environment will be turbulent, not only as a result of plasma disturbances caused by spacecraft passage and perturbation of the ambient density, but also because of such activities as thruster firings and water dumps. The turbulence could lead to the generation of electromagnetic emissions with a wide range of frequencies, leading in turn to possible problems with sensors, instrumentation, communications and control systems.

All these effects will have an impact upon future programmes in LEO and polar orbits. The possible phenomena that could occur, the possible consequences of the interactions with the spacecraft, and ways of monitoring and alleviating the effects require detailed study and analysis.

Spacecraft wake effects

The hypersonic motion of a spacecraft through the ionospheric plasma creates a shock wave in front and to the side of the body, and a large region behind the body which is depleted in ions. When the body size becomes large compared to the Debye length, then the interaction effects relative to the filling in of the wake become more severe. Typical wake lengths are many times the spacecraft body dimension.

The filling of the wake behind a large body moving supersonically or the expansion of a plasma into a rarefied area has been studied. Multiple charged particle populations result in polarisation electric fields which control particle motion along with flow expansion in the collisionless case and diffusion in cases where collisions must be considered. From these processes the wake region becomes a source of electron heating and ion acceleration, preferentially of lighter ions and of minor constituents. Other processes involve plasma oscillations and instabilities, strong jump discontinuities in plasma parameters at the expansion front, and rarefaction wave propagation into the ambient plasma. These phenomena all depend on the ionic constituents and concentrations, ambient electron temperature and density gradients, and the size of the body relative to the Debye length.

Spacecraft contamination effects

Early observations indicated that the Shuttle's local environment was controlled by the movement of the Shuttle through the ambient medium and by contaminant sources on the Shuttle. These sources, in the form of particulates and gases, are generated by Reaction Control System (RCS) and Orbital Manoeuvring System (OMS) engine firings, cabin gas leaks, water releases and outgassing of materials. Initial operational concerns over contamination focused on particulates scattering light into Shuttle based optical detectors to produce false signals, and on gaseous contaminants condensing on thermal control and optical sensing surfaces to degrade their performance.

Observations suggest that the Shuttle may be immersed in a large gas cloud, made of atoms and molecules from various outgassing sources, whose shape is governed by the Shuttle's interaction with the ambient neutral atmosphere and space plasma environment. Ionisation of contaminants and charge exchange ionisation processes lead to the formation of pick-up ions, tailing behind the spacecraft. Engine firings enhance the contaminant cloud and may produce their own characteristic contaminant cloud or plume that has an associated engine firing light flash which illuminates the Shuttle and enhances the surface glow phenomena. Particulate contamination is also enhanced when RCS engine exhaust plumes impinge directly on Shuttle surfaces. All of these observations suggest a close coupling between the various contaminant sources which contribute to the formation of a multispecious gas cloud surrounding the Shuttle.

Spacecraft power system effects

Exposed voltages on any part of a spacecraft cause current to flow between the element and the ambient low energy plasma environment. As an example, current flow to a solar array terminal could result in unacceptable power losses. At present, most spacecraft power systems use voltages less than 50 V. In future, it is planned to use much higher voltages. Although actual flight experience in space is limited, initial experiments have shown that leakage currents increase non-linearly with high positive voltages and that arcing occurs for high negative voltages. In addition, there are a number of other environmental effects on solar arrays in LEO that contribute to degradation and cause problems.

Spacecraft electromagnetic effects

Any AC and DC electric and magnetic fields on or near a spacecraft will be driven by two sources:
  1. spacecraft electromagnetic interference (EMI) associated with the hardware itself;
  2. fields associated with the interaction of the spacecraft and its environment.
In the case of the Space Shuttle, the environment was found to be dominated not by Orbiter generated noise but by plasma interaction noise.

Analyses of the wave environment aboard the Space Shuttle have led to the emerging picture that the broadband noise environment is dominated not by the induced environment associated with the large body interaction (although these effects certainly have a role to plasy), byt by the production of waves by the neutral gas cloud as it expands and undergoes chemical interactions, such as charge exchange which results in an ion tail, and creates plasma waves presumed to be driven by the ion currents. If this is the case, then it should be possible to correlate the level of background noise with the density of the neutral cloud, and indeed this has been done in the case of water releases from the Shuttle.

Evaluation of in-orbit spacecraft charging

Spacecraft charging effects

The buildup of large potentials on spacecraft relative to the ambient plasma is not, of itself, a serious electrostatic discharge (ESD) design concern. However, such charging enhances surface contamination, which degrades thermal properties. It also compromises scientific missions seeking to measure properties of the space environment. Spacecraft systems referenced to structure ground are not affected by a uniformly charged spacecraft. However, spacecraft surfaces are not uniform in their material properties, surfaces will be either shaded or sunlit, and the ambient fluxes may be anisotropic. These and other charging effects can produce potential differences between spacecraft surfaces or between spacecraft surfaces and spacecraft ground. When a breakdown threshold is exceeded, an electrostatic discharge can occur. The transient generated by this discharge can couple into the spacecraft electronics and cause upsets ranging from logic switching to complete system failure. Discharges can also cause long term degradation of exterior surface coatings and enhance contamination of surfaces. Vehicle torquing or wobble can also be produced when multiple discharges occur. The ultimate results are disruptions in spacecraft operation.

Surface charging could disrupt environmental measurements on scientific spacecraft. For this application and others where control of electrostatic fields is required, meterial selection to minimise differential charging is mandatory. For operational spacecraft, surface charging can also cause problems. The hallmark of the spaceraft charging phenomena is the occurrence of electronic switching anomalies. These anomalies are believed to result from transients caused differential charging induced discharges. These anomalous events even seem to occur in systems that are supposedly immune to noise. The discharge induced transients, under very severe environmental conditions, can cause system failures.

Conditions required for spacecraft charging

Work with the Defence Meteorological Satellite Program (DMSP) allows the required conditions for spacecraft charging (spacecraft surface potentials at least 100 V different from the potential of the surrounding plasma) to be defined. These conditions are:
  1. The spacecraft must encounter an auroral electron plasma which imposes on it a sufficiently large ratio of high energy electron ambient flux to total ion (ambient or ram flux).
  2. The ambient electron plasma has a relativel large fraction of its total flux at energies well above the secondary yield maximum of the spacecraft surface material, in order to suppress secondary emission effects which would tend to discharge the spacecraft.
  3. The spacecraft must be in darkness in order to suppress photoemission effects, which would tend to discharge the spacecraft.
    In addition, theoretical work suggests that:
  4. Charging is more likely for larger spacecraft because electron collection increases more rapidly with spacecraft size than ion collection does.

There are two aspects of the charging phenomena that should be kept in mind when assessing the impact upon future spacecraft systems:

  1. Recorded charging levels were for spacecraft with typical dimensions of one metre. As noted above, the charging levels are expected to increase in relation to an increase in the spacecraft dimension. Thus, while maximum charging levels of the order of -0.5 kV have been observed on polar transiting spacecraft to date, levels of charging in excess of -1 kV are anticipated for larger spacecraft. This must be considered during the design of such systems.
  2. Presumably, the charging levels recorded by the spacecraft instrumentation represented absolute charging levels, with the whole vehicle assuming the potential with respect to the plasma. Possible more serious in the context of future systems is the question of differential charging. Ion density depletion levels in the wake of a spacecraft, or components of a spacecraft, can be very large and very high levels of charge could accumulate upoon surfaces or equipment, if subject to high energy precipitating electrons. Again, this is an area that must be considered during the design of the systems.

Summary of previous work

The integrated study, and understanding, of a wide range of spacecraft/plasma interactions and their effect in low Earth and polar orbits must be an important component of efforts directed to the design, construction, and successful and safe operation of future spacecraft and large space systems.

Several of these problem areas have been investigated in laboratory experiments. However, it is important to appreciate that the simulation of a complex environment, such as the low Earth orbit ionosphere, in the laboratory is a difficult task. In general, there are five areas to which close attention should be paid:

A comprehensive review of test facilities used in experimental simulation studies led to the following conclusions:

The problems have also been studied using numerical and computational simulation methods. There are essentially two different philosophies that can be applied when writing computer codes to simulate the spacecraft/plasma interaction in low Earth orbit:

It is a feature of analytical codes that they describe only macro-processes (i.e. above the scale of individual particles) that are explicitly included by the author. Hence, in a sense, they tell the user only what he/she already knows and will not surprise him/her with unexpected behaviour.

There are generally less approximations in particle tracking codes than in analytical ones. Where approximations appear, they are, for example, in the form of the secondary electron emission spectrum and not in the particle trajectories.

In additon to this ground based work, account must be taken of space based experience. Over the past years, an ever increasing number of publications have appeared on the plasma modifications around a vehicle orbiting in the low Earth plasma. The mid-latitude, low Earth orbit region has been investigated fairly extensively by a series of satellites plus a few sounding rockets. In polar regions on the other hand, very few satellites have yet been flown which carried adequate instrumentation to monitor plasma modifications. However, many rocket payloads have been launched into the auroral zones.

Schematic of interactions with
    ambient plasma
Figure 2. Spacecraft interactions with ambient plasma

The current status of affairs can be summarised as follows:

The ESPIRE computer codes

Scope

The Spacecraft/Plasma Interactions and Electromagnetic Effects Program Suite ESPIRE was developed for ESA to be used as an aid to spacecraft design, addressing some specific problem areas in spacecraft/plasma interactions. This computer program suite was complemented by experimental work to aid the code validation. This was carried out in two different plasma simulation facilities. One had a large volume and a high gas pumping speed, and was capable of creating conditions which enabled large spacecraft with dimensions of hundreds of Debye lengths to be simulated. The second facility was smaller, but had the ability to vary or to cancel the magnetic field present in the plasma.

Experimental work on plasma wake phenomena used simple metallic models in the shape of a disc and a rectangular plate, placed in a streaming plasma. The wake structure downstream of the model was studied as a function of the Debye length or body size, model potential and Mach number. A valid simulation requires control of the slow ion population, using high pumping speeds and low operating pressures, and this was successfully achieved in both facilities.

Experimental studies in support of spacecraft charging used high energy (5-20 keV) electron guns to bombard the wake side of a model placed in a streaming plasma. Isolated metal targets of different materials and isolated Kapton were placed in the wake position. The charging levels were measured as a function of such parameters as ambient ion density, background pressure, ion energy, and test geometry.

The final outcome has been the development of a flexible software suite to model important aspects of spacecraft/plasma interactions and spacecraft charging. The emphasis has been upon experimentally validated engineering tools for use in spacecraft design. The suite is constructed in such a way as to facilitate future additions tot he programs as a result of the continuing development of engineering software tools.

Description of the ESPIRE codes

The interaction of a space vehicle with its surroundings in low Earth orbit is a complex phenomenon and no simple description will allow all the features to be adequately quantified. The use of computer codes does, however, provide the spacecraft designer with a powerful tool to analyse some of the effects since many of the physical processes occurring may be included in such coded. Nevertheless, despite the potential of computational methods, it is not realistic, at present, to design a code which includes all the physical processes that may be present in a rigorous and self-consistent manner. Indeed, such a code might be too slow or cumbersome to use for simple calculations of importance to the spacecraft designer.

The solution to this issue is the creation of a suite of computer programs each capable of analysing a part of the spacecraft/plasma interaction problem on the basis of a specific but restricted set of assumptions. While each code individually will allow only part of the problem to be addressed, the entire suite will allow a much more complete picture to be obtained. The ESPIRE program suite structure is illustrated in Fig. 3.

Structure chart of ESPIRE
Figure 3. Structure of the Spacecraft/Plasma Interactions and Electromagnetic Effects Program Suite ESPIRE

LEOPOLD

LEOPOLD Is a simple menu driven code whose function is the rapid determination of the principal parameters that characterise the low Earth orbit environment. The input to the code is the orbital altitude.

Using empirical and theoretical relationships and in-built data bases, the code evaluates and outputs seventeen space environment parameters. The input menus have been replaced by HTML forms in the SPENVIS implementation of LEOPOLD.

SOLARC

SOLARC Is the first of two programs lying at Level 2 in the software suite and as such is somewhat more complicated than LEOPOLD.

SOLARC Is an O-D code which provides an assessment of the current collection and the power loss that would be experienced by a solar array in LEO and polar environments. For a given set of conditions, specified by the user, the code creates current-voltage characteristics derived from two sets of empirical formulae by finding the solar array voltage distribution that produces a zero net collected current.

The first set of equations (model 1) were based upon the results of ground-based solar array experiments undertaken at the NASA Lewis Research Center. The second set of equations (model 2) were derived from space based data obtained on the PIX-2 test flight.

The equations are fully integrated with the SOLARC code. Any of the program variables (such as solar array area, voltage relative to space plasma potential, electron and ion current densities and temperatures) may be altered independently within the code.

EQUIPOT

The second code residing at Level 2 is EQUIPOT. EQUIPOT Is a flexible menu-driven code which permits a rapid assessment of the likelihood of charging for surface materials on a spacecraft. It utilises a simple geometry, a small isolated patch of material on a spherical spacecraft, calculates the various components of current to both the body (structure) and the patch, and estimates the equilibrium potentials which will develop in order to achieve zero net current. Approximate charging time is also computed. The user selects the structure and patch materials, and also the plasma environment by defining energy spectra for electron and ion fluxes. In the SPENVIS implementation of EQUIPOT, the menus have been replaced with HTML forms.

A number of options are available to cater for plasma regimes with Debye length large (GEO) or small (LEO), solar illumination or shadow, spacecraft velocity (ram and wake effects), and normal or isotropic incidence of particles. For small Debye lengths, it is reasonable to assume that current collection is sheath limited (plane probe assumption); for large Debye lengths, the geometry becomes more important and a spherical probe assumption offers an alternative limit.

SAPPHIRE

SAPPHIRE Is the first of two programs lying at Level 3 in the software suite. The programs at this level are specialised and require operator familiarity with the codes in order to run them. Due to their complexity, the Level 3 programs are outside the scope of SPENVIS, and have not been implemented.

SAPPHIRE Is a two dimensional particle and potential code which calculates the ion and electron densities and the electrostatic potential about a body or bodies in the presence of a streaming plasma. It has the capability of representing a variety of spacecraft geometries and allowing different potentials to be specified. A number of ambient plasma conditions may be modelled including monoenergetic plasmas and plasmas with a Maxwellian temperature distribution.

PICCHARGE

PICCHARGE Is particle-in-cell code that requires few non-physical assumptions. In the restricted situations to which it is limited, it produces accurate simulations of object/plasma interactions. Complex shapes can be modelled, with different materials making up their surfaces, and the charge buildup on different surface areas can be monitored. The object can be placed anywhere within the simulation space so that, with drifting plasmas, ram and wake effects can be examined.

The DICTAT computer code

Electrical charging of dielectric materials in the magnetosphere is a major cause of satellite anomalies. Where surface charging is concerned, there are a number of software tools (e.g. NASCAP (NASA Charging Analyzer Program) and EQUIPOT) which enable satellite designers to model the extent of the problem and to make satellites more resistant to this effect. For the internal charging problem a useful scientific tool is provided by the ESA-DDC code [Soubeyran and Floberhagen, 1994]. DICTAT Was developed to provide a practical engineering tool to address problems of internal dielectric charging.

DICTAT calculates the electron current that passes through a conductive shield and becomes deposited inside a dielectric. From the deposited current, the maximum electric field within the dielectric is found. This field is compared with the breakdown field for that dielectric to see if the material is at risk of an electrostatic discharge.

References

Martin, A. R., Spacecraft/Plasma Interactions and Electromagnetic Effects in LEO and Polar Orbits, Final Report for ESA/ESTEC Contract No. 7989/88/NL/PB(SC), Vol. 3, 1991.

, , Final Report for ESA/ESTEC Contract No. 7989/88/NL/PB(SC), Vol. 1, 199

Soubeyran, A., and R. Floberhagen, ESA-DDC 1.1 User Manual, Matra Marconi Space, 1994.


This text is based on Volume 3 of the Final Report for ESA/ESTEC Contract No. 7989/88/NL/PB(SC) (Martin, 1991).

Last update: Mon, 01 Mar 2010