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NASCAP calculations have been performed simulating exposure of a spacecraft-like model to multiple electron guns. The results agree well with experiment. It is found that magnetic field effects are fairly small, but substantial differential charging can result from electron gun placement. Conditions for surface flashover are readily achieved.
A computer code, the NASA Charging Analyzer Program (NASCAP), was developed to study the surface charging of bodies subjected to geomagnetic substorm conditions. This program will treat the material properties of a surface in a self-consistent manner and calculate the electric fields in space due to the surface charge. Trajectories of charged particles in this electric field can be computed to determine if these particles enhance surface contamination. A preliminary model of the Spacecraft Charging At The High Altitudes (SCATHA) satellite was developed in the NASCAP code and subjected to a geomagnetic substorm environment to investigate the possibility of electrostatic contamination. The results indicate that differential voltages will exist between the spacecraft ground surfaces and the insulator surfaces. The electric fields from this differential charging can enhance the contamination of spacecraft surfaces.
A jointly planned U.S. Air Force-NASA program has been established to investigate the spacecraft charging phenomenon that has caused electronic anomalies in satellites in geosynchronous orbits. The objectives of this program are to provide design criteria, techniques, and test methods to insure control of absolute and differential charging of spacecraft surfaces. These objectives will be updated continuously over the next four years as data become available from the combined contractual and in-house programs. The geosynchronous altitude environment will be defined, ground and flight tests will be conducted, and materials and charge control techniques will be developed as required. The ultimate output of the program will be a spacecraft charging design criteria and test specification document. The program will be coordinated by a spacecraft charging program review group which has both Air Force and NASA representation.
The role that energetic particles in the substorm plasma have on the charging and discharging of typical dielectric layers used on spacecraft was investigated using spectra and pitch angle distributions measured in situ on the SCATHA spacecraft prior to and during a few kilovolt differential charging event in eclipse conditions. The particle spectra was input to deposition codes that determine the dose rate as a function of depth in Kapton and Teflon layers used in an experiment on SCATHA. The calculated ambient dose rates of a few rads/sec throughout the bulk of the samples are sufficiently high that radiation damage levels are reached on the time scale of 1 year. Surface dose is a factor of 100 higher. Bulk conductivity profiles were obtained from the dose rates using empirical relationships. The radiation induced bulk conductivities calculated at the peak charging time are found to be smaller than the intrinsic dark conductivity range of solar conditioned Kapton but higher than the corresponding value for Teflon.
This paper considers the charging of the Space Shuttle Orbiter by energetic particles of environmental origin and from emission by accelerators. The results indicate that precipitating electrons quickly induce large voltages. High voltages may also occur when onboard accelerators inject energetic beams into the high altitude plasma. A significant conclusion from electron beam experiments is that the rockets charged to positive potentials much less than anticipated from the theory of probes in a quiescent plasma. Elementary theories predict the large negative potentials observed by firing energetic ions and predict severe differential charging of the Orbiter.
The environmental charging of satellite surfaces during geomagnetic substorms is the apparent cause of a significant number of anomalous events occurring on geosynchronous satellites since the early 1970's. Electromagnetic pulses produced in connection with the differential charging of insulators can couple into the spacecraft harness and cause electronic switching anomalies. An investigation conducted to determine the response of the spacecraft surfaces to substorm particle fluxes makes use of a harness transient detector. The harness transient detector, called the Transient Event Counter (TEC) was built and integrated into the Canadian-American Communications Technology Satellite (CTS). A description of the TEC and its operational characteristics is given and the obtained data are discussed. The data show that the satellite surfaces appear to be charged to the point that discharges occur and that the discharge-induced transients couple into the wire harnesses.
The ion engine on ATS-6 has been operated in daylight and eclipse. The effect on particle fluxes to the spacecraft was monitored. This data provides information on the potential of the spacecraft with respect to the ambient plasma and the local electric fields caused by the charge distribution on the spacecraft. Operation of the ion engine and/or the neutralizer clamps the spacecraft within a few volts of the ambient plasma potential. Operation of only the neutralizer tends to reduce the differential charge on the satellite, whereas operation of the ion engine reduces it below the sensitivity limit of the detectors.
A two dimensional model was developed to describe the charging of thin polymer films exposed to a uniform mon-energetic electron beam. The study was motivated by observed anomalous behavior of geosynchronous satellites which was attributed to electrical discharges associated with the differential charging of satellite surfaces of magnetospheric electrons. Electric fields both internal and external to the irradiated specimen were calculated at steady state in order to identify regions of high electrical stress. Particular emphasis was placed on evaluating the charging characteristics near the material's edge. The model was used to identify and quantify the effects of some of the experimental parameters notably: beam energy; beam angle of incidence; beam current density; material thickness; and material width. Simulations of the following situations were also conducted: positive or negative precharging over part of the surface; a central gap in the material; and a discontinuity in the material's thickness.
Plasma data from the University of California, San Diego, Auroral Particles experiments on Applied Technology satellites 5 and 6 are used to investigate the dynamics of natural charging events. Both eclipse and daylight charging events are considered, and typical responses are illustrated by data from specific events. Two different physical processes are found to be involved in the charging process. One of these is straightforward: the spacecraft structure potential responds rapidly to changes in the environment, typically changing by hundreds of volts in a few seconds. The other process is more subtle: differential charging and potential barrier formation precede structural charging and determine the time scale; typically, tens of minutes are required for the potential to change by several hundreds of volts. The latter process is found to be predominately responsible for daylight charging on both spacecraft.
A detailed model of the geometrical, material, and electrical properties of the SCATHA satellite for use with the NASA charging analyzer program is described. Charging calculations in an intense magnetospheric substorm environment demonstrated that: (1) long booms can significantly perturb the potentials near the spacecraft, and (2) discharging by sunlight or by active control can cause serious time-dependent differential charging problems.
A two-dimensional model is developed to describe the charging of strips of thin polymer films above a grounded substrate exposed to a uniform mono-energetic electron beam. The study is motivated by the observed anomalous behavior of geosynchronous satellites, which has been attributed to differential charging of the satellite surfaces exposed to magnetospheric electrons. Surface and bulk electric fields are calcuated at steady state in order to identify regions of high electrical stress, with emphasis on behavior near the material's edge. The model is used to study the effects of some of the experimental parameters, notably beam energy, beam angle of incidence, beam current density, material thickness and material width. Also examined are the consequences of a central gap in the material and a discontinuity in the material thickness.
The ion engine experiments on ATS 6 were operated in daylight and eclipse. The effect on particle fluxes to the spacecraft was monitored with the UCSD Auroral Particles Experiment. These data also provide information on the potential of the spacecraft with respect to the ambient plasma and on the local electric fields caused by the charge distribution on the satellite. Daylight operations of the plasma bridge neutralizer and the cesium thruster in fall, 1974, served to hold the spacecraft between -3 and -8 volts with respect to the ambient plasma. Neutralizer operation reduced differential charging effects, while operation of the thruster usually reduced the effects below the detectors sensitivity. Eclipse operations of the neutralizer reduced kilovolt negative potentials to a few volts. Operation of the thruster prevented possible charging of the satellite during substorms, making it possible to study low energy particle spectra which are at times obscured by charging during substorms.
Techniques are discussed for controlling the potential of a spacecraft by means of devices which release appropriate charged particles from the spacecraft to the environment. Attention is given to electron emitters, ion emitters, a basic electron emitter arrangement, techniques for sensing electric field or potential, and flight experiments on active potential control. It is recommended to avoid differential charging on spacecraft surfaces because it can severely affect the efficacy of emitters. Discharging the frame of a spacecraft with dielectric surfaces involves the risk of stressing the dielectric material excessively. The spacecraft should, therefore, be provided with grounded conductive surfaces. It is pointed out that particles released by control systems can return to the spacecraft.
Studies of the Earth with the ATS-5, ATS-6, and SCATHA spacecraft led to the development of several simple tools for predicting the potentials to be expected on a spacecraft in the space environment. These tools have been used to estimate the expected levels of worst case charging at Jupiter and Saturn for the Galileo and the Cassini spacecraft missions. This paper reviews those results and puts them in the context of the design issues addressed by each mission including the spacecraft design mitigation strategies adopted to limit differential charging. The model shows that shadowed surfaces in Earth orbit can reach 25 kV or higher in worst case environments. For Galileo, spacecraft- to-space potentials of 900 V were predicted in shadow. Since such potentials could produce possible discharges and could effect low energy plasma measurements, the outer surface of Galileo was designed to rigid conductivity requirements. Even though the surface of Galileo is not entirely conducting, after 27 orbits no adverse effects due to surface charging aside from limited effects on low energy plasma measurements have been reported. The saturnian environment results in spacecraft potentials to space in shadow of 100 V or less. Although the overall surface of the Cassini spacecraft was not entirely conducting and grounded, it is shown that only in the most extreme conditions, is it expected that Cassini will experience any effects of surface charging at Saturn.
The environment near the surface of asteroids, comets, and the Moon is electrically charged due to the Sun’s photoelectric bombardment and lofting dust, which follows the Sun illumination as the body spins. Charged dust is ever present, in the form of dusty plasma, even at high altitudes, following the solar illumination. If a body with high surface resistivity is exposed to the solar wind and solar radiation, sunexposed areas and shadowed areas become differentially charged. The E-Glider (Electrostatic Glider) is an enabling capability for operation at airless bodies, a solution applicable to many types of in-situ missions, which leverages the natural environment. This platform directly addresses the "All Access Mobility" Challenge, one of the NASA’s Space Technology Grand Challenges. Exploration of comets, asteroids, moons and planetary bodies is limited by mobility on those bodies. The lack of an atmosphere, the low gravity levels, and the unknown surface soil properties pose a very difficult challenge for all forms of know locomotion at airless bodies. This EGlider levitates by extending thin, charged, appendages, which are also articulated to direct the levitation force in the most convenient direction for propulsion and maneuvering. The charging is maintained through continuous charge emission. It lands, wherever it is most convenient, by retracting the appendages or by firing a cold-gas thruster, or by deploying an anchor. Preliminary calculations indicate that a 1 kg mass can be electrostatically levitated in a microgravity field with a 2 m diameter electrostatically inflated ribbon structure at 19kV, hence the need for a “balloon-like” system. The wings could be made of very thin Au-coated Mylar film, which are electrostatically inflated, and would provide the lift due to electrostatic repulsion with the naturally charged asteroid surface. Since the E-glider would follow the Sun’s illumination, the solar panels on the vehicle would constantly charge a battery. Further articulation at the root of the lateral strands or inflated membrane wings, would generate a component of lift depending on the articulation angle, hence a selective maneuvering capability which, to all effects, would lead to electrostatic (rather than aerodynamic) flight.
The Microwave Anisotropy Probe (MAP), a MIDEX mission built in partnership between Princeton University and the NASA Goddard Space Flight Center (GSFC), will study the cosmic microwave background. It will be inserted into a highly elliptical earth orbit for several weeks and then use a lunar gravity assist to orbit around the second Lagrangian point (L2), 1.5 million kilometers, anti-sunward from the earth. The charging environment for the phasing loops and at L2 was evaluated. There is a limited set of data for L2; the GEOTAIL spacecraft measured relatively low spacecraft potentials (approx. 50 V maximum) near L2. The main area of concern for charging on the MAP spacecraft is the well-established threat posed by the "geosynchronous region" between 6-10 Re. The launch in the autumn of 2000 will coincide with the falling of the solar maximum, a period when the likelihood of a substorm is higher than usual. The likelihood of a substorm at that time has been roughly estimated to be on the order of 20% for a typical MAP mission profile. Because of the possibility of spacecraft charging, a requirement for conductive spacecraft surfaces was established early in the program. Subsequent NASCAP/GEO analyses for the MAP spacecraft demonstrated that a significant portion of the sunlit surface (solar cell cover glass and sunshade) could have nonconductive surfaces without significantly raising differential charging. The need for conductive materials on surfaces continually in eclipse has also been reinforced by NASCAP analyses.
The NASA Charging Analyzer Program (NASCAP) has been validated in a space environment. Data collected by the SCATHA (Spacecraft Charging at High Altitude) spacecraft has been used with NASCAP to simulate the charging response of the spacecraft ground conductor and dielectric surfaces with considerable success. Charging of the spacecraft ground observed in eclipse, during moderate and severe substorm environments, and in sunlight has been reproduced using the code. Close agreement between both the currents and potentials measured by the SSPM's, and the NASCAP simulated response, has been obtained for differential charging. It is concluded that NASCAP is able to predict spacecraft charging behavior in a space environment.
The sensitivity of the charging response to the representation of the measured environments and material properties are discussed. Single and double Maxwellian representations are compared with direct numerical integration of the observed spectra. The effect of anisotropic incident flux distribution is modeled. In addition, the effect of the high energy radiation upon bulk conductivity and hence differential charging is examined.