Todd A. Schneider

Electrostatic Discharge Testing of Multijunction Solar Array Coupons After Combined Space Environmental Exposures

A set of multijunction GaAs/Ge solar array test coupons provided by Space Systems/Loral was subjected to a sequence of five-year increments of combined space environmental exposure tests. The test coupons capture an integrated design intended for use in a geosynchronous (GEO) space environment. A key component of this test campaign is performing electrostatic discharge (ESD) tests in the inverted gradient mode. The protocol of the ESD tests is based on the ISO standard for ESD testing on solar array panels [ISO-11221]. The test schematic in the ISO reference has been modified with Space System/Loral designed circuitry to better simulate the on-orbit operational conditions of its solar array design. Part of the modified circuitry is to simulate a solar array panel coverglass flashover discharge. All solar array coupons used in the test campaign consist of four cells constructed to form two strings. The ESD tests are performed at the beginning-of-life (BOL) and at each five-year environmental exposure point. The space environmental exposure sequence consists of ultraviolet radiation, electron/proton particle radiation, thermal cycling, and xenon ion thruster plume erosion. This paper discusses the coverglass flashover simulation, the ESD test setup, and the importance of the electrical test design in simulating the on-orbit operational conditions. Results from fifth-year testing are compared to the baseline ESD characteristics determined at the BOL condition.

DIRECT DRIVE HALL THRUSTER SYSTEM DEVELOPMENT

The status of development of a Direct Drive Hall Thruster System is presented. In the first part, a study of the impacts to spacecraft systems and mass benefits of a direct-drive architecture is reviewed. The study initially examines four cases of SPT-100 and BPT-4000 Hall thrusters used for north-south station keeping on an EXPRESS-like geosynchronous spacecraft and for primary propulsion for a Deep Space-1 based science spacecraft. The study has also been extended to include the impact of direct drive on orbit raising for higher power geosynchronous spacecraft and on other deep space missions as a function of power and velocity change. The major system considerations for accommodating a direct drive Hall thruster are discussed, including array regulation, system grounding, distribution of power to the spacecraft bus, and interactions between current-voltage characteristics for the arrays and thrusters. The mass benefit analysis shows that, for the initial cases, up to 42 kg of mass savings is attributable directly to changes in the propulsion hardware. When projected mass impacts of operating the arrays and the electric power system at 300V are included, up to 63 kg is saved for the four initial cases. Adoption of high voltage lithium ion battery technology is projected to further improve these savings. Orbit raising of higher powered geosynchronous spacecraft is the mission for which direct drive provides the most benefit, allowing higher efficiency electric orbit raising to be accomplished in a limited period of time, as well as nearly eliminating significant power processing heat rejection mass. The total increase in useful payload to orbit ranges up to 278 kg (11%) for a 25 kW spacecraft, launched from an Atlas IIA. For deep space missions, direct drive is found to be most applicable to higher power missions with a velocity change up to several km/s, typical of several Discovery-class missions. In the second part, the status of development of direct drive propulsion power electronics is presented. The core of this hardware is the heater-keeper-magnet supply being qualified for the BPT- 4000 by Aerojet. A breadboard propulsion power unit is in fabrication and is scheduled for delivery late in 2003.

Charging of the International Space Station as Observed by the Floating Potential Measurement Unit: Initial Results

The floating potential measurement unit (FPMU) is a multiprobe package designed to measure the floating potential of the International Space Station (ISS) as well as the density and temperature of the local ionospheric plasma environment. The purpose of the FPMU is to provide direct measurements of ISS spacecraft charging as continuing construction leads to dramatic changes in ISS size and configuration. FPMU data are used for refinement and validation of the ISS spacecraft charging models used to evaluate the severity and frequency of occurrence of ISS charging hazards. The FPMU data and the models are also used to evaluate the effectiveness of proposed hazard controls. The FPMU consists of four probes: a floating potential probe, two Langmuir probes, and a plasma impedance probe. These probes measure the floating potential of the ISS, plasma density, and electron temperature. Redundant measurements using different probes support data validation by interprobe comparisons. The FPMU was installed by ISS crew members during an extra-vehicular activity on the starboard (S1) truss of the ISS in early August 2006 when the ISS configuration included only one 160-V U.S. photovoltaic (PV) array module. The first data campaign began a few hours after installation and continued for over five days. Additional data campaigns were completed in 2007 after a second 160-V U.S. PV array module was added to the ISS. This paper discusses the general operational characteristics of the FPMU as integrated on ISS, the functional performance of each probe, the charging behavior of the ISS before and after the addition of a second 160-V U.S. PV array module, and initial results from model comparisons.

Magnetic filter type plasma source for ground-based simulation of low earth orbit environment

Simulation of the low Earth orbit (LEO) plasma environment in ground-based vacuum facilities is important for studies of spacecraft interaction with ionospheric plasmas. In this paper we describe the design and performance of a magnetic filter-equipped plasma source. Experimental data collected in the expanding plasma downstream of the source suggest it is a good candidate for use as a LEO plasma simulator in that the expanding plasma has a very low electron temperature and contains streaming ions—the plasma environment encountered by satellites in LEO. Adjustable plasma source operating conditions of flow rate, discharge current and discharge voltage enable production of plasma electron temperatures over the range from 0.17 to 0.35 eV and streaming ion energies over the range from 1 to 4 eV.

Combined Space Environmental Exposure Test of Multijunction GaAs/Ge Solar Array Coupons

The purpose of this test program is to understand the changes and degradation of the space solar array panel components, including its electrostatic discharge (ESD) mitigation design features in their integrated form, after multiple years (up to 15) of simulated Earth geosynchronous (GEO) space environment. A set of multijunction GaAs/Ge solar array test coupons was subjected to five-year increments of combined environmental exposure tests. These tests consisted of the following: simulated ultraviolet (UV) radiation, ESD, electron/proton particle radiation, thermal cycling, and simulated ion thruster exposures. The solar radiation simulation was produced using a mercury-xenon lamp with wavelengths in the UV spectrum ranging from 230 to 400 nm. The ESD test was performed in the inverted-gradient mode using a low-energy electron (3-6 keV) beam exposure. The ESD test also included a simulated panel coverglass flashover for the primary arc event. The electron/proton radiation exposure included 1.0-MeV electrons, 100-keV electrons, and 40-keV protons. Thermal cycling included simulated transient Earth eclipse for satellites in geosynchronous orbit. With the increasing use of ion thruster engines on many satellites, the combined environmental exposure test also included ion thruster exposure to determine the impact on solar array performance, as well as the ion thruster interaction to ESD events. Before and after each increment of combined environmental exposures, the coupons underwent visual inspection using high power magnification and electrical tests that included characterization by large-area pulse solar simulator, dark I -V, insulation resistance, and electroluminescence. This paper discusses the test objective, test methodologies, and preliminary results after five and ten years of simulated combined environmental exposure tests.

First Preliminary Results From U.S. Round-Robin Tests

The first preliminary results are reported from the U.S. Round-Robin Test on Plasma Expansion Speed. The tests were performed at the NASA Glenn Research Center on two coupons (six strings) of International Space Station (ISS) solar cell arrays, with a separate small array to obtain arcs (because it is so difficult to get ISS arrays to arc). ISS arrays were used because they have no exposed interconnects to act as bare current collectors and confuse the experimental results. The preconstructed ISS strings were laid out approximately parallel to the plasma expansion velocity to allow for the best test of the simple plasma expansion front current waveform model. Several Langmuir probes were arranged above and to the sides of the sample to allow for measurement of the plasma propagation speed. In the initial set of tests, primary arcs and the consequent current waveforms were measured in a Low Earth Orbit-type plasma. In a second set of tests, two electron guns with diffusers were used to provide an approximately uniform Geosynchronous Earth Orbit-type environment, and primary arcs and current waveforms were obtained. The objective of this and other round-robin tests is to characterize primary arc waveforms in terms of speed and degree of discharge of arc plasmas produced by primary arcs, and their dependences on environment, capacitance per unit area, arc voltage, temperature, and so on. The final goal is to allow engineering estimates of arc current peaks and half-widths to allow confident design and construction of space solar arrays, and to allow mitigation techniques to be evaluated. This is the first in an extended series of tests to be performed at six different U.S. facilities, and with participation from ten different U.S. organizations.

Survey of International Space Station Charging Events

With the negative grounding of the 160V Photovoltaic (PV) arrays, the International Space Station (ISS) can experience varied and interesting charging events. Since August 2006, there has been a multi-probe p ackage, called the Floating Potential Measurement Unit (FPMU), availa ble to provide redundant measurements of the floating potential of th e ISS as well as the density and temperature of the local plasma environment. The FPMU has been operated during intermittent data campaigns since August 2006 and has collected over 160 days of information reg arding the charging of the ISS as it has progressed in configuration from one to three PV arrays and with various additional modules such as the European Space Agency?s Columbus laboratory and the Japan Aeros pace Exploration Agencys Kibo laboratory. This paper summarizes the charging of the ISS and the local environmental conditions that contr ibute to those charging events, both as measured by the FPMU.

Electrostatic Discharge Tests on Solar Array Wire Coupons Subjected to Simulated Space Environment Aging

Solar array wire coupons have successfully completed an environmental life test that simulates 15 years at geosynchronous orbit. The environments included: ultraviolet (UV), electron, and proton irradiation; thermal cycling; electrostatic discharge (ESD); and ion thruster plume exposure on an electrostatically charged coupon. The test articles consisted of two wire coupons: one simulated the sun-facing side and the other simulated the shade side. UV irradiation was performed only on the sun-facing side coupon. Tests were performed at beginning-of-life, 7.5 years, and end-of-life. At each age point, ESD tests were performed using an electron beam at a worst case geosynchronous space environment flux of 1 nA/cm2. The test setup included capacitance that simulated a whole panel array harness, and a solar array simulator that generated in-flight array current profiles. The test coupon configuration contained aspects of the full panel wiring topology, which included potential fault conditions on wires. Visual inspections, documented with photographs, and isolation resistance tests were performed after each environment exposure to ensure the integrity of the wire insulation. This paper/presentation discusses each environment test level, test condition, and results from the various environmental age points.

Validation of the Plasma Densities and Temperatures From the ISS Floating Potential Measurement Unit

The validation of the floating potential measurement unit (FPMU) plasma density and temperature measurements is an important step in the process of evaluating International Space Station (ISS) spacecraft charging issues including vehicle arcing and hazards to crew during extravehicular activities. The highest potentials observed on the Space Station are due to the combined Vsp times B effects on a large spacecraft and the collection of ionospheric electron and ion currents by the 160-V U.S. solar array modules. The ionospheric plasma environment is needed for input to the ISS spacecraft charging models used to predict the severity and frequency of occurrence of ISS charging hazards. The validation of these charging models requires the comparison of their predictions with measured FPMU values. The FPMU measurements themselves must also be validated for use in manned flight safety work. This paper presents preliminary results from a comparison of densities and temperatures derived from the FPMU Langmuir probes and plasma impedance probe with the independent density and temperature measurements from a spaceborne ultraviolet imager, a ground-based incoherent scatter radar, and ionosonde sites.

Solar Arrays for Direct-Drive Electric Propulsion: Electron Collection at High Voltages

Solar array technologies that can empower electric thrusters in direct drive mode may provide significant mission benefits by reducing power processing, system complexity, weight, and cost over conventional systems. In direct-drive systems the solar arrays will operate at high voltages and must perform safely in the surrounding plasma environment, with minimal loss of performance due to parasitic current collection. For example, current state-of-the-art Hail effect thrusters in the kilowatt class require applied voltages of 300 V or higher. Results from experiments and modeling of electron current collection at high bias voltages (300-500 V) are presented. The experiments employed two sample solar array coupon technologies. A hollow cathode was used to emulate the induced environment around the solar arrays far from the Hall thruster and in nearby regions populated by charge-exchange plasma (10 1 2 -10 1 3 m - 3 and 0.5-1 eV). The measurements show that tens to hundreds of seconds are required before the collected current relaxes to a quasi-steady value. Comparisons with results from numerical calculations suggest that changes of the secondary electron yield properties of the dielectric materials may account for the observed current collection trends.