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Touchless Potential Sensing of Complex Differentially-Charged Shapes Using Secondary Electrons
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[The final version of this work can be found at http://doi.org/10.2514/1.A35355] The secondary electron method has been recently proposed to touchlessly sense the electrostatic potential of non-cooperative objects in geosynchronous equatorial orbits and deep space. This process relies on the detection of secondaries generated at the target surface, that is irradiated by an electron beam. Although the concept has been demonstrated with basic geometries, the electric field around a complex body leads to a highly inhomogeneous distribution of secondary electrons that determines the performance of the system. This paper employs vacuum chamber experiments and particle tracing simulations to investigate the detectability of the secondary electron flux generated over a spacecraft-like electrode assembly. The differential charging scenario, in which the assembly is charged to multiple potentials, is also studied. A three-dimensional particle tracing framework that implements the coupled electron beam propagation and secondary electron generation processes is introduced and validated, showing its utility as a diagnostic tool. The spacecraft shape, potential distribution, and electron beam intersection define the detectability of the target, which is limited to well-defined spatial regions where the potentials are measured with high accuracy. The analysis provides theoretical and technical insight into the development of future electron-based remote potential sensing technologies.
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... Novel active sensing methods have been recently proposed to touchlessly sense the electrostatic potential of non-cooperative objects in Geosynchronous Equatorial Orbit (GEO) and deep space. Such approaches make use of a positively charged servicing craft that directs an electron beam at the object of interest so that low-energy secondary electrons [1,2] and x-rays [3,4] are emitted from its surface. The servicer measures the incoming signals and, knowing its own potential, infers that of the target. ...
... However, the physics of each problem are not favorable to the simultaneous generation of these signals: while secondary electrons are produced at moderate electron beam energies [16], the generation of x-rays is favored by energetic particle impacts [3]. In addition, low-energy electron beams are steered in the presence of the inhomogeneous electrostatic field generated by the servicer-target system, increasing the uncertainty of the problem [2,15]. From a technical perspective, it would be convenient to develop a sensing procedure that uncouples both mechanisms and optimizes the generation and control of secondary electrons and x-rays while minimizing the current fluxes imparted on the target. ...
... In addition to δ max and E max , the secondary electron flux also depends on the backscattered electron yield η(E, ϕ). Even though the degradation of these parameters with respect to laboratory conditions can significantly impact the charge balance and secondary electron flux magnitude in the target, the spatial detectability of secondaries in a complex electrostatic environment is generally considered to remain unaffected [2]. If further assurance is desired, the use of a collimator may be used to limit the sources of secondary electrons entering the detector and reject SEE emissions from background surfaces (see e.g. the Electron Drift Experiment for the Magnetospheric Multiscale Mission [46]). ...
Ultraviolet lasers are proposed as a replacement for low-energy electron beams to induce the emission of secondary electrons in touchless spacecraft potential sensing technologies. Theoretical considerations show that the measurement process becomes significantly less sensitive to the electrostatic environment and leads to more robust, controllable systems. Lasers could be employed in combination with high-energy electron beams to independently induce the emission of photoelectrons and x-rays close to their optimum operational points. This approach would enable hybrid photoelectron and x-ray potential sensing methods with enhanced detectability and sensing accuracy. Applications in touchless potential sensing, charge control, and material characterization are identified.
... Both components are made of aluminum. Additionally, a Retarding Potential Analyzer (RPA) is included in the setup and used to touchlessly estimate potentials with the electron method [20], but is not required for the x-ray method. The electron beam is a EMG-4212C from Kimball Physics and capable of emitting electrons with energies from 1-30 keV and currents from 1 A to 100 A. The focus of the electron beam is adjustable, which allows to either bombard a large area of the target object with electrons, or to focus the electron beam on a small spot. ...
... SIMION solves Laplace's equation to derive the electric field and then computes the particle trajectory from Newton's second law. The implementation of the SIMION simulation framework for remote sensing of electric potentials is discussed in greater detail in Reference [20]. Figure 3 shows the SIMION model of the experimental setup. ...
... For electrode configuration (a) and an angle of 40 • -50 • , the x-ray detector measures the potential of the panel even though the SIMION simulation predicts the electron beam to hit both the bus and the panel. However, small modeling inaccuracies of the experimental setup geometry can have a large effect on the accuracy of the SIMION simulation [20]. Thus, this discrepancy could be explained by an imprecise SIMION chamber model. ...
A method has been recently proposed to estimate the electric potential of co-orbiting spacecraft remotely using x-rays that are excited by an electron beam. Prior work focused on the theoretical foundation and experimental validation of this method using flat target plates. Although useful for the validation of this concept, flat plates do not adequately represent the shape of spacecraft and the resulting complex particle dynamics. Additionally, all previous experiments were conducted with fully conducting test objects, but components of spacecraft are not always connected to one common electric ground. This paper experimentally investigates the remote electric potential estimation of objects with complex shapes and differentially-charged components using x-rays. A particle tracing simulation software is used to assist the interpretation of the experimental results. The results show that the orientation of the target strongly affects which component's potential is measured. A new analysis method is proposed that enables the simultaneous measurement of multiple potentials using a single x-ray spectrum.
... N OVEL active and passive potential sensing of neighboring spacecraft has been investigated in recent years and deemed feasible for application the Geosynchronous region [1][2][3][4]. Active potential sensing involves a servicing spacecraft directing an electron beam at a target so that secondary electrons [5,6] and x-rays [7,8] are emitted from the surface, as shown in Figure 1. The use of a low-wavelength ultraviolet laser has also been investigated as a means to excite photoemissions from a target [9,10]. ...
... For an eclipse environment, which may occur if the target is shadowed by the servicer or other spacecraft, the secondary electron current is the dominant emitted current. The angular distribution of secondary electrons approximately follows a Lambertian distribution and is nearly independent of the impacting particle's angle of incidence [6,48]. Therefore, the range of angles is defined as zero to π, or the highest range of emitted angles possible. ...
Touchless potential sensing of neighboring spacecraft using photoemissions and secondary electron emissions has been investigated for geosynchronous (GEO) applications. As more missions are being sent to cislunar space, this technology may be extended there as well. However, the complexity of the cislunar environment presents novel challenges for touchless potential sensing technology. A chief issue is shorter Debye lengths than in GEO regions, which can be as low as 10 m in the cislunar regions. Therefore, a model for the electric and potential fields surrounding a charged spacecraft in short Debye regions around the moon is investigated. The vacuum (Laplace) and Debye–Hückel models are presented and effective Debye lengths are used to expand the models and better represent the environment. The effective Debye length has previously been investigated in low Earth orbit (LEO), quiet GEO, and asteroid environments but has not been found in the cislunar plasma environment, and a larger effective Debye length may allow touchless potential sensing using electron emissions to be possible at farther, safer distances than expected. Once the effective Debye lengths and associated models are established, the relationship between effective Debye lengths and touchless potential sensing capabilities is explored through computations in NASCAP-2k, a spacecraft–plasma interaction software. The developed methods are then used to determine whether passive and active touchless potential sensing is feasible in cislunar regions with nonnegligible electrostatic potential shielding.
A method has been recently proposed to estimate the electric potential of co-orbiting spacecraft remotely using x-rays that are excited by an electron beam. Prior work focused on the theoretical foundation and experimental validation of this method using flat target plates. Although useful for the validation of this concept, flat plates do not adequately represent the shape of spacecraft and the resulting complex particle dynamics. Additionally, all previous experiments were conducted with fully conducting test objects, but components of spacecraft are not always connected to one common electric ground. This paper experimentally investigates the remote electric potential estimation of objects with complex shapes and differentially-charged components using x-rays. A particle tracing simulation software is used to assist the interpretation of the experimental results. The results show that the orientation of the target strongly affects which component's potential is measured. A new analysis method is proposed that enables the simultaneous measurement of multiple potentials using a single x-ray spectrum.
Novel active sensing methods have been recently proposed to measure the electrostatic potential of non-cooperative objects in geosynchronous equatorial orbit and deep space. Such approaches make use of electron beams to excite the emission of secondary electrons and X-Rays and infer properties of the emitting surface. However, the detectability of secondary electrons is severely complicated in the presence of complex charged bodies, making computationally efficient simulation frameworks necessary for in-situ potential estimation. The purpose of this paper is twofold: firstly, to introduce and test a quasi-analytical, uncoupled, and computationally efficient electron beam expansion and deflection model for active charging applications; and secondly, to characterize the uncertainty in the beam-target intersection properties, which condition the measurement of secondary electrons. The results show that a combination of secondary electrons and X-ray methods is highly desirable to yield a robust and accurate measure of the potential of a target spacecraft.
The electrostatic tractor has been proposed to touchlessly remove space debris from geosynchronous orbit by taking advantage of intercraft Coulomb forces. A controlled spacecraft (tug) emits an electron beam onto an uncooperative or retired satellite (debris). Thus, the tug raises its own electrostatic positive potential to tens of kilovolts, whereas the debris charges negatively. This results in an attractive force called the electrostatic tractor. Prior research investigated the charged relative motion dynamics and control of the electrostatic tractor for two spherical spacecraft and how charge uncertainty affects the relative motion control stability, but attitude effects could not be studied due to the two-sphere model. This work uses the multisphere method to consider general three-dimensional spacecraft shapes, and it investigates how the electric potential uncertainty and debris attitude impact the equilibrium separation distance between the two craft. The results show bounds for safe operations that avoid a collision. State regions are identified where the relative motion is particularly sensitive to potential uncertainty. The relative station-keeping performance using either higher- or lower-fidelity multi-sphere method models are compared to demonstrate that even a lower-fidelity multi-sphere method model can yield good results.
Remote surface potential determination for nearby objects in space is an enabling technology for a range of space mission concepts, from improved rendezvous capabilities to touchless debris remediation with the electrostatic tractor concept. One concept is to use a nearby servicing spacecraft which fires electrons at the target. These electrons generate bremsstrahlung X‐rays, and the resultant X‐ray spectrum can be used to determine the landing energy of the electrons. By knowing the initial energy of the electrons, such as those emitted from an electron gun, the relative potential of the target can be inferred. The use of electron‐induced X‐rays to determine the electrostatic potential of a surface has previously been investigated theoretically, but this work demonstrates it experimentally. Accuracy is found to correlate to the angle between the detector and the target, and investigation of this relationship is a topic of future work. The mean error in landing energy is found to be approximately 1%, and the mean error in estimated plate voltage is found to be within the uncertainty of the high voltage power supply. Material determination is successfully demonstrated, using characteristic X‐rays to identify elements that make up less than 0.5% by mass of the target sample. Several possible means of improving landing energy accuracy are discussed.
Remote charge sensing is a technique that can provide valuable insight into spacecraft‐environment interactions, as well as enable missions that leverage electrostatic interactions between multiple spacecraft or involve docking maneuvers in harsh charging environments. The concept discussed in this paper uses a co‐orbiting servicing spacecraft to measure the energies of secondary electrons or photoelectrons that are emitted from the target object with initial energies of a few electron volts. The electrons are accelerated toward the servicing spacecraft, which is at a known positive potential (relative to the target), where they are measured by an energy analyzer. Given the potential of the servicing spacecraft, the potential of the target can be accurately determined. Results are presented from experiments conducted in a vacuum chamber to investigate the touchless sensing concept. Specifically, the feasibility and accuracy of the electron method is considered for different materials, charge scenarios, and relative geometries. The results show that the surface potential of a flat plate can be accurately determined for a range of metallic surface materials, voltages, and relative angles.
As spacecraft does not have an independent method to conduct charge to ground, it naturally accumulates charge due to interactions with the ambient plasma and surface emission. This charge produces an electric field surrounding the spacecraft, which takes the form of a plasma sheath. Charged particles traveling through this sheath are altered in both energy and direction, thus affecting derived scientific quantities. While this effect has been known since the advent of space based particle instruments, this work represents the first time that an in situ characterization study of this effect has been possible. The Fast Plasma Investigation, of the Magnetospheric Multiscale mission, obtains near simultaneous measurements of phase space, via particle counts from 512 look directions and 32 energies. These new data allow the relative effects of the plasma sheath to be explored at a high time and spatial resolution. This work presents a method by which these measurements are used to study a ground model that traces the migration of particles through the sheath and estimates the error in measured velocity. This approach, performed statistically, leads to an estimation of uncertainty in particle count distributions for a given location in phase space and characterization of the redistribution of counts within a skymap due to sheath effects.
Secondary electron emission yield from the surface of SiC ceramics induced by Xe¹⁷⁺ ions has been measured as a function of target temperature and incident energy. In the temperature range of 463–659 K, the total yield gradually decreases with increasing target temperature. The decrease is about 57% for 3.2 MeV Xe¹⁷⁺ impact, and about 62% for 4.0 MeV Xe¹⁷⁺ impact, which is much larger than the decrease observed previously for ion impact at low charged states. The yield dependence on the temperature is discussed in terms of work function, because both kinetic electron emission and potential electron emission are influenced by work function. In addition, our experimental data show that the total electron yield gradually increases with the kinetic energy of projectile, when the target is at a constant temperature higher than room temperature. This result can be explained by electronic stopping power which plays an important role in kinetic electron emission.
Sheaths are formed around the surfaces of airless bodies, which both collect and emit electrons. Depending on the ratio of the emitted to collected electron fluxes Γ, the sheath potential structure can be quite different. We present the first experimental measurements of all the three types of the sheath potentials: classical, space-charge-limited (SCL), and inverse. A solid surface immersed in plasmas emits secondary electrons (SEs). The potential structure changes from a monotonic classical to a non-monotonic SCL sheath as Γ increases. At the critical electron emission with zero electric field at the surface, the sheath potential is determined by the plasma electron temperature, and Γ approaches but remains smaller than 1. The non-monotonic SCL sheath persists steadily for Γ > 1. When the emitted electron density becomes larger than the plasma electron density, a monotonic inverse sheath forms with a positive surface potential relative to the ambient.
In this paper, the electrostatic sheath of a simplified spacecraft is investigated for heliocentric distances varying from 0.044 to 1 AU, using the 3-D Particle in Cell software Satellite–Plasma Interaction System. The baseline context is the prediction of sheath effects on solar wind measurements for various missions, including the Solar Probe Plus mission (perihelion at 0.044 AU from the sun) and Solar Orbiter (SO) (perihelion at 0.28 AU). The electrostatic sheath and the spacecraft potential could interfere with the low-energy (a few tens of eV) plasma measurements, by biasing the particle distribution functions measured by the detectors. If the spacecraft charges to large negative potentials, the problem will be more severe as low-energy electrons will not be seen at all. The Solar Probe Plus and SO cases will be presented in details and extended to other distances through a parametric study, to investigate the influence of the heliocentric distance to spacecraft. Our main result is that, for our spacecraft model, the floating potential is a few volts positive from 1 AU to about 0.3 AU, while below 0.3 AU, the space charge of the photoelectrons and secondary electrons create a potential barrier that drives the spacecraft potential negative.
A three-dimensional (3-D), self-consistent code is employed to solve for the static potential structure surrounding a spacecraft in a high photoelectron environment. The numerical solutions show that, under certain conditions, a spacecraft can take on a negative potential in spite of strong photoelectron currents. The negative potential is due to an electrostatic barrier near the surface of the spacecraft that can reflect a large fraction of the photoelectron flux back to the spacecraft. This electrostatic barrier forms if (1) the photoelectron density at the surface of the spacecraft greatly exceeds the ambient plasma density, (2) the spacecraft size is significantly larger than local Debye length of the photoelectrons, and (3) the thermal electron energy is much larger than the characteristic energy of the escaping photoelectrons. All of these conditions are present near the Sun. The numerical solutions also show that the spacecraft's negative potential can be amplified by an ion wake. The negative potential of the ion wake prevents secondary electrons from escaping the part of spacecraft in contact with the wake. These findings may be important for future spacecraft missions that go nearer to the Sun, such as Solar Orbiter and Solar Probe Plus. Comment: 25 pages, 7 figures, accepted for publication in Physics of Plasmas
Secondary electron emission of surfaces exposed to oscillating electromagnetic field is at the origin of the multipacting effect that could severely perturb the operation of particle accelerators. This contribution tries to illustrate by measurement results, the origin of the secondary electron emission as well as the main reasons for the discrepancies between technical materials and pure metals. The variation of the secondary electron yield with the incident electron energy will be discussed for various types of technical surfaces. The influence of a gas condensation on these surfaces will also be addressed in the context of the LHC accelerator. Various treatments aiming at a permanent reduction of the secondary electron yield will be presented. A special attention will be paid to the decrease of the secondary electron yield under electron or photon impact and to its possible beneficial consequences for the processing of devices prone to multipacting.
Plasma and field measurements on board magnetospheric spacecraft can be influenced by the spacecraft potential and emitted particles such as photoelectrons and secondary electrons. There exist contradictory requirements for, on one hand, minimizing the spacecraft potential and, on the other hand, minimizing perturbations in the sheath and contamination by spacecraft-generated electrons. Assessment and mitigation of such effects therefore require improved quantitative modeling of the spacecraft electrostatic sheath. In this paper a fully self-consistent model of the plasma around an electron-emitting central body in a spherically symmetric geometry is used to analyze the electrostatic sheath around an idealized magnetospheric spacecraft. Although the model is too simplistic to allow detailed comparisons with observations, it helps to analyze the phenomenon of potential barrier occurrence and some global characteristics, which can be relevant to conductive magnetospheric spacecraft like Geotail or Cluster. It is shown that nonmonotonic potential with negative potential barrier can exist all around a positively charged spacecraft even in the case of realistic illumination pattern. The magnitude of the potential barrier in regions of the magnetosphere with large Debye-length plasma is found to be as large as a few volts negative for small positive spacecraft potential but that it quickly vanishes when the potential is increased by a few volts. Furthermore, the location of the potential barrier is found to be much more variable than previously predicted.
Differential charging effects observed in the electron data of the University of California, San Diego, auroral particles experiment on Applied Technology Satellite 6 are described and analyzed. An electrostatic barrier around the environmental measurements experiment (EME) package on ATS 6 is shown to be the natural result of dielectrics around the spacecraft which are more negatively charged than the mainframe of the spacecraft. In particular, the large dish antenna on ATS 6 causes the formation of a barrier which traps particles emitted from the EME package surface and returns them to the spacecraft. The insulating surface of the rotating University of Minnesota detector on the otherwise conducting EME package is shown to be the source of accelerated fluxes of electrons during charging events.
Artificial objects currently populating Earth orbital regimes can be distinguished by comparing remote observational data to that of optical measurements within the visible (VIS) and near infrared (NIR) spectral regions of materials obtained in the laboratory. Careful comparison of spectrally resolved observations with laboratory-based material irradiation experiments can provide a tool for remote diagnosis and anomaly resolution. However, the spectral signatures of many materials change as a function of exposure to radiation in geosynchronous Earth orbit (GEO). In order for remote characterization to be a viable characterization technique, the evolution of material optical properties (VIS-NIR) upon exposure to the GEO environment must be understood. To this end, we have studied the reflectance of several commonly used spacecraft materials including organic polymers, solar cell coverglasses, black thermal control paint, and carbon-carbon (C–C) matrix composite as a function of exposure to high energy electron irradiation in order to mimic the harsh environment of GEO. It was found that different classes of materials exhibit large variation in radiation stability resulting in the development of unique spectral characteristics.
Numerous missions are being proposed which involve multiple spacecraft operating in close proximity in harsh charging environments. In such missions, the ability to sense the electrostatic potential on a nearby object is critical to prevent harmful electrostatic discharges or to leverage Coulomb interactions for relative motion control. The electron method is one such technique for touchless potential measurement which works by measuring low-energy secondary or photoelectrons emitted from the target. Previous work has demonstrated the efficacy of the electron method for touchless sensing, but has been limited to consideration of simple shapes and uniformly charged targets. This paper investigates the electron method for touchless sensing for cases in which the target spacecraft has more complex geometry primitives, including boxes, panels, and dishes. Further, the differential charging case, in which the target object is charged to multiple, different potentials, is also considered. A simulation framework is developed to model electric fields and particle trajectories around such spacecraft geometries. Vacuum chamber experiments validate the simulation results. The study shows how the target geometry can focus or defocus the electron flux into streams of electrons emanating from the surface. This provides critical insight into where to place the servicer vehicle to measure these fluxes and determine the target spacecraft potential.
High tumbling rates of uncooperative target pose strong technical challenges and collision risks that can prevent removal of the debris using contact, such as with robotic arms or capture nets. Electrostatic touchless detumbling is a promising technology that can be used to decrease the rotational velocity of an uncooperative object in geosynchronous orbit, from a safe distance. This paper demonstrates the advantages of applying a Lyapunov optimal control in conjunction with a surface multisphere model. This approach allows for the analysis of general shapes, eliminating the need for analytical approximations on debris shape and expected torque, employed by previous work. Moreover, using this model, the robustness of the system to uncertainties to the debris center of mass position is tested. This analysis uncovers an unstable phenomenon that was previously not captured using simpler models. An active disturbance rejection control ensures robustness of the system in the cases analyzed, also granting an increase in its effectiveness. It is shown in simulation that the system can exploit deviations in the center of mass to achieve a higher level of controllability and completely detumble all components of angular velocity.
[The final version of this work can be found at https://doi.org/10.1016/j.actaastro.2021.12.037] The Electrostatic Charging Laboratory for Interactions between Plasma and Spacecraft (ECLIPS) research vacuum chamber facility has recently been developed as part of the Autonomous Vehicle Systems Laboratory at the University of Colorado Boulder. This experimental spacecraft charging research facility provides capability to conduct experiments relevant to charged astrodynamics in a space-like environment. The paper discusses the development, characterization, and present capabilities of the vacuum chamber, which includes a range of sources to provide electron, ion, and photon fluxes, a 3-axis motion stage, magnetic field control, and a variety of sensors. This state-of-the art facility has been used to conduct experiments on touchlessly sensing spacecraft potentials, electrostatic actuation experiments, development of a new type of electron gun, and will continue to be a unique facility for studying spacecraft electrostatics in the future.
Although spacecraft charging is often thought of as a purely harmful phenomenon, if controlled it can be used as a means of touchless actuation. If a tug spacecraft irradiates a nearby debris object with an electron beam, the tug will charge positively and the debris will charge negatively. This creates an attractive Coulomb force that the tug can use to touchlessly tug the debris object from the geosynchronous orbit into a graveyard orbit. Compared with earlier work this paper uses a more advanced charging model with isotropic fluxes for the calculation of electron and ioninduced yields, and an empirical model of electron and ion fluxes rather than Maxwellian distributions. This new model is used to calculate the attractive force for a variety of tug to debris size ratios, beam currents and voltages, and the inclusion of pulsing the electron beam. The major result of using this new charging model is that it takes more current, and thus power, than was used in prior work to charge a debris object due to the higher yields from isotropic fluxes. The electrostatic tractor concept can still move a range of large, tumbling debris objects to the graveyard orbit in a few months.
This paper presents the Global Positioning System (GPS) satellites arc into the space environment. Evidence for this consists of undispersed impulsive radio frequency signals are shown in the dual-frequency RF receivers, that are on-board all GPS satellites, with high rates that track the charging electron environment, detection of appropriate amplitude short-timescale pulses from GPS satellites at 327 and 430 MHz with the Arecibo 305 -m Gordon radiotelescope that do not appear in off-source scans, and arcing on laboratory GPS-like solar arrays at voltages similar to those inferred from GPS arc thresholds. In addition, GPS satellites (up to the Block IIRM) have power degradation (in excess of that expected from radiation) of about 1.5% per year. This degradation is primarily in the short-circuit current, consistent with contamination on the solar array cover glasses, which was detected by a thermal monitor on one GPS satellite solar array, and consistent with the power degradation observed on a highly arced GPS-like laboratory solar array if extrapolated to GPS undispersed event rates. The GPS arc rate required for solar array contamination and consistent with the onboard observed event rates are very high, in fact, higher than the rates measured previously on SCATHA, CTS, and other satellites. It is conjectured that the effects of these arcs on satellite power systems have been routinely prevented by highly filtering the power system and antenna response, that most satellites have had no way to detect arcing, and that most power losses up to the present time have been ascribed to radiation damage. Results of new optical and radio-frequency radio frequency observations of satellite arcing are discussed, as well as laboratory measurements confirming the high arc rates and their consequences.
Measuring the charge on a nearby space object during close proximity, servicing, and rendezvous and docking operations without requiring physical touch remains challenging despite decades worth of data on spacecraft charging and its risks. This paper proposes the means to identify the charge on a closely neighboring space object and its elemental composition by examining the X-ray spectrum generated by energetic electrons impacting the target. In particular, deconvolution of the bremsstrahlung X-ray continuum provides an estimate of the landing energy of the electrons. Knowing the initial electron energy, the potential difference between the source and the target is determined. Additionally, characteristic X-rays emitted during the process of energetic electron-matter impact allows the relative abundance of elements in the target to be determined. Spatial separations in the order of tens of meters are required between an electron gun and the corresponding detector to maximize the collection of the bremsstrahlung spectrum. This could be achieved with either the single-craft and deployable booms or through the use of two spacecraft. Electron beam energies of 40 kV are found to generate sufficient levels of X-rays for potential determination at over 10 m from the target and to determine the landing energy of the beam to within 0.14% using commercial X-ray detectors.
A method is investigated to use secondary electrons to touchlessly sense the electrostatic potential of an object in geosynchronous orbit or deep space. This method involves a positively charged servicing craft, which directs a high-energy electron beam at the object of interest such that low-energy secondary electrons are emitted from the surface. The low-energy electrons emitted by the target are accelerated toward the servicing craft and arrive with an energy equal to the potential difference between the two crafts. The servicing craft measures the electron energy spectrum and, knowing its own potential, then infers the potential of the target. Depending on the application, photoelectrons could similarly be used to infer the target potential without the need for an active electron beam. Though it is possible to measure potential by directly contacting a surface, remote measurement offers significant advantages and supports missions which must operate in close proximity without making physical contact. Several missions have been proposed that use interactions between charged objects to create useful forces and torques, including electrostatic detumbling and reorbiting of debris, Coulomb formations, and virtual structures. Remote measurement of potential would benefit these missions by enabling feedback control of the active charging. Other applications include mitigation of arcing during rendezvous, docking, and proximity operations for future servicing or salvaging missions.
The number of operational satellites and debris objects in the valuable geosynchronous ring has increased steadily over time such that active debris removal missions are necessary to ensure long-term stability. These objects are very large and tumbling, making any mission scenarios requiring physical contact very challenging. In the last 10 years, the concept of using an electrostatic tractor has been investigated extensively. With the electrostatic tractor concept, active charge emission is employed to simultaneously charge the tug or services vehicle, while aiming the charge exhaust onto the passive space debris object to charge it as well. The resulting electrostatic force has been explored to actuate this debris object to a disposal orbit or to detumble the object, all without physical contact. This paper provides a survey of the related research and reviews the charging concepts, the associated electrostatic force and torque modeling, and the feedback control developments, as well as the charge sensing research.
The plasma electrons bombarding a plasma-facing wall surface can induce secondary electron emission
(SEE) from the wall. A strong SEE can enhance the power losses by reducing the wall sheath
potential and thereby increasing the electron flux from the plasma to the wall. The use of the materials
with surface roughness and the engineered materials with surface architecture is known to reduce
the effective SEE by trapping the secondary electrons. In this work, we demonstrate a 65% reduction
of SEE yield using a velvet material consisting of high aspect ratio carbon fibers. The measurements
of SEE yield for different velvet samples using the electron beam in vacuum demonstrate the dependence
of the SEE yield on the fiber length and the packing density, which is strongly affected by the
alignment of long velvet fibers with respect to the electron beam impinging on the velvet sample.
The results of SEE measurements support the previous observations of the reduced SEE measured
in Hall thrusters. Published by AIP Publishing. https://doi.org/10.1063/1.4993979
This chapter describes processes involved in immunizing a spacecraft against spacecraft charging problems, and provides design guidelines. The system developer should demonstrate through design practices, test, and analysis that spacecraft charging effects will not cause a failure to meet mission objectives. Analysis should be used to evaluate a design for charging in the specified orbital environment. Testing usually ranks high among the choices to verify and validate the survivability of spacecraft hardware in a given environment. Some of the test discussed here include: material testing, component testing, assembly testing and system testing. The chapter provides general guidelines and quantitative recommendations on design guidelines/techniques that should be followed in hardening spacecraft systems to spacecraft charging effects. The design guidelines are divided into subsections: general, surface charging, internal charging, solar arrays, and special situations. aerospace materials; building integrated photovoltaics; surface charging
The geosynchronous (GEO) orbital regime is becoming cluttered with derelict space debris, which raises the collision risk for satellites in a very valuable orbit. Touchless reorbiting options have been proposed for moving this debris to a graveyard orbit to avoid the risk of physically docking with large, multiton defunct satellites that can be tumbling at several degrees per second. This paper investigates the electrostatic tractor (ET), which uses an electron beam mounted on a tug spacecraft to irradiate a passive debris object. The tug quickly rises to a positive steady-state voltage, and the debris quickly falls to a negative steady-state voltage to create an attractive tugging force. The tug maintains a fixed relative position during the reorbit using inertial thrusting. The Coulomb force can be used as a means of touchless actuation for geosynchronous debris reorbiting, detumbling, and formation flying. This paper investigates the prospects of using a beam with pulsed current. The off-pulsing periods can have benefits for sensing and thrusting applications, and the pulsing charging can lead to higher force levels for the same electrical power used in particular conditions. A Monte Carlo analysis is performed to study the mean electrostatic force, considering a range of beam currents, voltages, pulsing duties cycles, and vehicle sizes. Power-limited regions are identified where the pulsed tractor has a magnitude that is comparable or greater than the mean continuous beam force. This creates interesting alternate methods to implement an ET while having periodic off periods. The tractor performance is illustrated through the debris reorbiting scenario. A detailed equal power analysis determines that even duties cycles as low as 10%-20% can lead to forces comparable to the continuous beam performance, or even do better for some powers.
As has been revealed in a number of more recent astrophysical papers, in most of the tenuous space plasmas Maxwellian distribution functions cannot be expected for ions or electrons because of the lack of efficient relaxation processes. Many of the classical characteristics of plasmas, such as plasma frequency or Debye length, are calculated on the basis of the assumption, however, that Maxwellians prevail, which under most of the relevant astrophysical plasma conditions is not the case. We here therefore consider this specific problem of Debye shieldings of single charges in a plasma for the case of prevailing non-equilibrium distribution functions for ions and electrons. As typical non-equilibrium functions, so-called Kappa functions were considered with clear preference, and we therefore study here the Debye shielding in a plasma with Kappa-distributed electrons and ions. We show that the so-called Debye shielding increases with increasing extent of the high-velocity tail of the electron distribution function, or in other words, with lower Kappa index of the underlying Kappa function. In our calculations we demonstrate that the Debye lengths become enlarged by about a factor of 10 with respect to its classically expexted value if highly suprathermal electron distributions prevail with Kappa indices close to 1.5.
Coulomb formation flight is a concept that utilizes electrostatic forces to control the separations of close proximity spacecraft. The Coulomb force between charged bodies is a product of their size, separation, potential and interaction with the local plasma environment. A fast and accurate analytic method of capturing the interaction of a charged body in a plasma is shown. The Debye-Hückel analytic model of the electrostatic field about a charged sphere in a plasma is expanded to analytically compute the forces. This model is fitted to numerical simulations with representative geosynchronous and low Earth orbit (GEO and LEO) plasma environments using an effective Debye length. This effective Debye length, which more accurately captures the charge partial shielding, can be up to 7 times larger at GEO, and as great as 100 times larger at LEO. The force between a sphere and point charge is accurately captured with the effective Debye length, as opposed to the electron Debye length solutions that have errors exceeding 50%. One notable finding is that the effective Debye lengths in LEO plasmas about a charged body are increased from centimeters to meters. This is a promising outcome, as the reduced shielding at increased potentials provides sufficient force levels for operating the electrostatically inflated membrane structures concept at these dense plasma altitudes.
[1] Double-probe electric field sensors installed on scientific spacecraft are often deployed using wire booms with radii much less than typical Debye lengths of magnetospheric plasmas (millimeters compared to tens of meters). However, in tenuous and cold-streaming plasmas seen in the polar cap and lobe regions, the wire booms, electrically grounded at the spacecraft, have a high positive potential due to photoelectron emission and can strongly scatter approaching ions. Consequently, an electrostatic wake formed behind the spacecraft is further enhanced by the presence of the wire booms. We reproduce this process for the case of the Cluster satellite by performing plasma particle-in-cell (PIC) simulations, which include the effects of both the spacecraft body and the wire booms in a simultaneous manner. The simulations reveal that the effective thickness of the booms for the Cluster Electric Field and Wave (EFW) instrument is magnified from its real diameter (2.2mm) to several meters, when the spacecraft potential is at tens of volts. Such booms enhance the wake electric field magnitude by a factor of 1.5–2 depending on the spacecraft potential and play a principal role in explaining the in situ Cluster EFW data showing sinusoidal spurious electric fields with about 10mV/m amplitude. The boom effects are quantified by comparing PIC simulations with and without wire booms and also by examining the wake formation for various spacecraft potentials.
This paper explores the concept of using electrostatic forces for deployment of gossamer space structures. The Electrostatically Inflated Membrane Structure (EIMS) uses two conducting membranes that are interconnected through membrane ribs. An absolute electrostatic charge is applied to the structure through active charge emission. This causes repulsion between layers of lightweight membranes that inflates the EIMS system and tensions the membranes. Assuming positive tensions, the EIMS system is modeled as a rigid system. Typical orbital perturbations are considered such as solar radiation pressure, differential gravity, and atmospheric drag which may compress the structure leading to shape destabilization. Restricting the analysis in this paper to flat membranes, the minimum potentials required to exactly compensate for the worst case scenario of differential solar radiation pressure at geostationary altitudes are estimated to be on the order of hundreds of volts. In low Earth orbit, voltage magnitudes of several kilovolts are required to reach an inflation pressure to offset the normal compressive drag pressure.
Using the three-dimensional, low-energy electron spectrometer aboard the Ulysses spacecraft, the authors have measured the gyrotropicity of electron distributions in the solar wind. In order to make these observations, they have developed a new technique for correcting spacecraft charging effects in three-dimensional, low-energy particle measurements. Comparisons of ion and electron number and current densities, and the alignment of electron temperature anisotropies with the local magnetic field, are presented as evidence of the improvement in the accuracy of the electron moments resulting from the spacecraft charging corrections. The implications of these charging correction technique go beyond simple scalar corrections to the Ulysses measurements. They discuss the effects of their charging correction upon the measurements of temporal and radial gradients in a plasma environment and for two-dimensionally obtained low-energy particle data. 17 refs., 12 figs.
The goal of this second edition is siniflar to the first: to allow you to begin with a ``blank sheet of paper`` and design a space mission to meet a set of broad, often poorly defined, objectives. You should be able to define the mission in sufficient detail to identify principal drivers and make a preliminary assessment of overan performance, size, cost, and risk. The emphasis of the book is on low-Earth orbit, unmanned spacecraft. However, we hope that the principles are broad enough to be applicable to other n-dssions as well. We intend the book to be a practical guide, rather than a theoretical treatise. As much as possible, we have provided rules of thumb, empirical formulas, and design algorithms based on past experience. We assume that the reader has a general knowledge of physics, math, and basic engineering, but is not necessarily familiar with any aspect of space technology. This book was written by a group of over 50 senior space engineers. It reflects the insight gained from this practical experience, and suggests how things might be done better in the future. From time to time the views of authors and editors conflict, as must necessarily occur given the broad diversity of experience. We believe it is important to reflect this diversity rather than suppress the opinions of individual authors. Similarly, the level of treatment varies among topics, depending both on the issues each author feels is critical and our overan assessment of the level of detail in each topic that is important to the preliminary mission analysis and design process. The book is appropriate as a textbook for either introductory graduate or advanced undergraduate courses, or as a reference for those already working in space technology.
In the Penning trap, there is transport of electrons in the limit of zero gas pressure that arises from asymmetric stray electric fields. In an annular version of the Penning trap, this asymmetry transport is shown to be greatly reduced when the plasma-facing surfaces are coated with colloidal graphite. In a separate device, an emissive probe is used to examine the space potential a few millimeters above coated and uncoated surfaces. It is found that the rms potential variation is approximately 250 mV for uncoated surfaces and 15 mV for coated surfaces. The characteristic length scale of the inhomogeneities is ~1 cm. Glow-discharge cleaning, which is easily renewed, is shown to reduce the potential variation to the same level that is obtained with the colloidal graphite coating.
The secondary electron emission from alkaline-earth oxide-coated cathodes has been investigated under both continuous and pulsed bombardment. Experiments have been performed with three types of apparatus. Yield vs. energy data reveal values of delta of 4-7 at room temperature, with a more or less flat maximum at approximately 1000 volts primary energy. The yield increases with temperature in an exponential manner, and plots of Deltadelta (i.e.,deltaK°- delta300°K) vs. 1T give straight lines. Values of Q1 between 0.9-1.5 ev are generally indicated, and from extrapolation of these curves, yields exceeding 100 at 850°C are deduced. The secondary emission depends upon the degree of activation, and increases with enhancement of the thermionic emission characteristics. Short time effects such as growth or decay of secondary current after the onset of primary bombardment or persistence after the cessation of bombardment have not been observed, and values of yield obtained by pulsed methods are in accord with those obtained under d.c. conditions. Tail phenomena reported by J. B. Johnson and interpreted as "enhanced thermionic emission" from oxide-coated cathodes become manifest only under experimental conditions characterized by certain space-charge effects, and have been effectively simulated by bombarding a tantalum target adjacent to an electron-emitting tungsten filament. Various measurements of the energy distribution of secondary electrons as a function of primary voltage and temperature have been obtained. It was observed that the average energy of the secondary electrons decreases with temperature at a rate which more than compensates for the increase in the number of secondaries emitted per incident primary. The mechanism of the observed dependence of yield upon temperature is not well understood. Various alternative explanations are discussed and, in the light of the present state of our knowledge, regarded as untenable.
To compute the ion current density at any point of a current-collecting
device mounted on a spacecraft, such as an electrostatic probe or the
aperture of a mass spectrometer, one must determine the properties of
all possible trajectories passing through that point. In the
steady-state collisionless case, this knowledge is sufficient if one
also knows the unperturbed ion velocity distribution at large distances,
because the phase-space density does not change along the trajectories.
We compute the ion current for a specific class of spacecraft
experiments. Included in this class are mass spectrometers with
attractive apertures, and ion traps (flush-mounted circular planar
probes with internal grids). The investigation is based on the
computation of trajectories in the electric fields due to spacecraft
potential, the drawing-in potential of the experiment aperture, and the
space charge. We consider two values of the Debye length, namely, an
infinite length (Laplace field) and a length comparable to the
experiment dimensions (but small compared with the spacecraft
dimensions). The Laplace field is calculated by solution of linear
difference equations in the region surrounding the spacecraft and is
independent of the ion velocity distribution. The small-Debye-length
field is estimated by a linearized approximation that also leads to
linear difference equations. We consider H+, He+,
and O+ ion currents. Either an attractive satellite potential
or an attractive drawing-in potential enhances the current by a large
factor. Another effect of the satellite potential is to reduce the
amplitude of the current modulation caused by the spacecraft spin. For
large Debye lengths and no drawing-in potential, the current is found to
depend in a simple manner on a parameter that is the ratio of the work
done on the ion by the electric field to the unperturbed ion kinetic
energy in the spacecraft reference frame. The so-called `planar
approximation' is poor for the Laplace field (or large Debye lengths)
but tends to improve as the Debye length is reduced. The current-voltage
characteristic of an internal repelling collector in the case of an ion
trap with no drawing-in potential is also investigated. It is found that
the temperature can be inferred from the shape of the retarded
current-voltage characteristic at sufficiently large retarding
potentials, regardless of the Debye length. Concentrations that have
been inferred either with large drawing-in potentials or under
large-Debye-length conditions, without corrections for the electric
field effects, are probably unreliable.
The change in height of the potential energy barrier at the surface of a metal with the expansion of the metal due to heating is investigated. Also, the change of the normal maximum energy of an electron in a metal is calculated as a function of the temperature of the metal. These calculations show that the work function of a metal is a linear function of its temperature. These results, when combined with the thermionic emission equation, show that the thermionic emission constant is a characteristic of the metal, and is no longer the same for all pure metals. The calculated results for the thermionic emission constants for several metals show fair agreement with the experimentally determined values.
Recently, several uses for intercraft Coulomb forces have been explored. Proposed applications have ranged from creating static formations of many spacecraft to steerable nanosat deployment systems. This paper considers the use of Coulomb forces for creating space structures. Unlike conventional space structures, these "virtual space structures" have no physical connections. Instead, they are held together by maintaining specific charges at their node points. This form of structure is thus readily expandable, can be reconfigured to different shapes, and is easily deployed. Fundamental concepts are covered in conjunction with a control law illustrating the ability to form a structure with conventional stiffness and damping coefficients that can be actively modified.
Spacecraft formation flying using Coulomb forces is a relatively new technology for spacecraft control, and may have application for a wide variety of mission objectives including attitude control, collision avoidance, and orbit perturbation correction. Coulomb-controlled formations appear ideally suited for close formation flying in high Earth orbits to perform wide field of view imaging missions using separated spacecraft interferometry. Preliminary research has shown that it is possible to induce Coulomb forces and torques comparable with those of candidate high-efficiency electric thrusters with less than 1 Watt of on-board power. This paper discusses the challenges and prospects of developing spacecraft formations utilizing Coulomb forces. Formation flying on the order of tens of meters is very difficult using conventional ion propulsion methods, because the exhaust plumes will quickly interfere with the delicate on-board sensors. The Coulomb forces would allow the relative motion of satellites to be controlled without such contaminations. The fuel efficiency of the control makes very long duration missions possible. Static equilibrium solutions of Coulomb formations are discussed. Further, the behavior of a two-satellite Coulomb formation is discussed at GEO. A nonlinear, orbit elements based control law is introduced to stabilize the relative motion and drive the orbit element differences toward desired values.
Measurements have been made of electron backscattering in a range of incident beam energy from 05 to 10 keV. The variation of backscattering coefficient ? with energy and angle of incidence, and the energy spectrum of backscattered electrons at incidence and take-off angles of 45?, have been measured. Results from carbon, aluminium-silicon, copper and silver were obtained at a pressure of approximately 10?? Torr to make the effects of contamination under the electron beam negligible.
Various simple theories of the backscattering phenomenon which have been used successfully at higher energies are examined in an attempt to understand the processes producing the observed results. Although the theory of elastic scattering of electrons by atoms at low energies is complicated, it is found that a simple theory based on largeangle elastic collisions may explain the experimentally observed rise in ? with fall in beam energy, when an accurate range-energy relationship is used. The variation of ? with angle of beam incidence can be explained for low atomic number elements on a diffusion theory, but this explanation does not wholly explain the results for copper at low energies. The energy spectra cannot be predicted by any existing simple theory.
Spacecraft in tenuous plasmas become positively charged because of photoelectron emission. If the plasma is supersonically drifting with respect to the spacecraft, a wake forms behind it. When the kinetic energy of the positive ions in the plasma is not sufficient to overcome the electrostatic barrier of the spacecraft potential, they scatter on the potential structure from the spacecraft rather than get absorbed or scattered by the spacecraft body. For tenuous plasmas with Debye lengths much exceeding the spacecraft size, the potential structure extends far from the spacecraft, and consequently in this case the wake is of transverse dimensions much larger than the spacecraft. This enhanced wake formation process is demonstrated by theoretical analysis and computer simulations. Comparison to observations from the Cluster satellites shows good agreement. (c) 2006 American Institute of Physics.
This paper gives an overview of electrostatic charging which occurs on spacecraft in different plasma environments. Particular emphasis is given to differential charging between sunlit and shadowed insulated surfaces, a phenomenon which is often observed in the geostationary orbit. It can generate potential differences of several kilovolts between adjacent surfaces. This can lead to discharges and serious spacecraft anomalies such as spurious telecommands caused by voltage and current transients on cable harnesses. Experience with the GEOS and ISEE satellites has demonstrated that differential charging can be avoided by making outer surface elements conductive and connecting them to a common ground.
A simple theoretical calculation is given in which the energy distribution of low‐energy secondary electrons emitted from metals is derived. The main feature of the calculation is an energy‐dependent electron mean free path. The theory predicts that the peak of the energy distribution occurs at a value of the secondary electron energy equal to φ/3, where φ is the metal work function.
High electrostatic potentials of satellite surfaces can result in electrostatic discharges (ESDs) that can damage or interfere with the satellite electronics. Measurements of the satellite conducting frame potential relative to the plasma environment are typically used to provide information on the statistical probability of charging. The frame potential is usually measured from the spectrum of the ions incident to the satellite or by voltage probes outside the satellite plasma sheath. However, ESD events are more directly related to the differential charging of various surface materials on the satellite. We will present satellite differential charging observations of several sample materials from the Satellite Surface Potential Monitor (SSPM) on the Spacecraft Charging at High Altitudes (SCATHA) mission. The SSPM data are compared statistically with measurements of the satellite frame potential and with the ESD events detected by the SCATHA discharge monitor. Local time and radial occurrence of frame and SSPM sample potentials are shown to be very similar. However, the functional relationship between the sample and frame potentials appears to differ widely from event to event.
Computer modelling of vacuum systems is often performed using the Monte-Carlo technique. One of the assumptions frequently made is that molecules originating at the boundaries of the system have a “cosine distribution”. The nature of this distribution is occasionally misinterpreted and consequently, the molecular distribution is physically incorrect.This article describes the main point of confusion by firstly stating the cosine law and describing one possible method of “generating” the correct distribution. It goes on to describe the distribution that results from the most common misinterpretation of the cosine law. This is illustrated by quantifying the effect on the modelled conductance for circular cylinders of various lengths.
The conditions under which multiple valued solutions occur by computing the floating potential of an isolated eclipses surface on a geosynchronous orbit spacecraft were examined. Different approximations for the electron spectra during a geomagnetic substorm were used. The result indicates that if the incident electron flux has a Maxwellian energy distribution, the ratio of the secondary emitted current to the incident electron current is independent of the spacecraft potential. In this case a single value solution to the current equation occurs.
SIMION (R) 8.1 User Manual, Rev-5
- D Manura
- D Dahl
Manura, D., and Dahl, D., SIMION (R) 8.1 User Manual, Rev-5, Adaptas Solutions, LLC, Palmer, MA 01069, 2008. URL
http://simion.com/manual/.
Physics and Applications of Secondary Electron Emission
- H Bruining
Bruining, H., Physics and Applications of Secondary Electron Emission, New York: McGraw-Hill Book Co., Inc. London:
Pergamon Press Ltd., 1954. Chapter 7: Theory of Secondary Electron Emission; Discussion of Some Properties of Secondary
Electrons.