Kim K. de Groh

An Overview of the Results from the 30,000 Hr Life Test of Deep Space 1 Flight Spare Ion Engine

The extended life test of the Deep Space 1 (DS1) spare flight ion thruster (FT2) was voluntarily terminated on June 26th 2003. Although the engine had not yet reached its end of life at the conclusion of the test, the decision to terminate and begin the post-test destructive analyses was made to benefit near term ion engine development programs. During its 5-year run, the thruster operated for a total of 30,352 hours, processed 235.1 kg of Xenon propellant, and acquired several thousand hours of operation at each of the four independent throttled conditions investigated. The objectives of the test were to characterize previously observed failure modes, identify unknown failure modes, and quantify thruster performance as a function of engine wear and throttle level. Several performance variations and degradation processes were observed and monitored during the course of the test. Degradation processes included erosion of the discharge cathode keeper, ion optics grid sputter erosion, and deposition of material in the neutralizer cathode at low power that later cleared with the return to full power operation. Performance degradation was limited to reduction in measured thrust at the full power point for the final 1000 hours of operation, most likely due to electronbackstreaming. Post-test inspection of the engine was initiated immediately following the test termination, to ascertain causes of the wear, and to look for any previously unknown wear processes. The ion engine consists of various internal systems, and the post-test analysis effort has been divided into separate analysis efforts of the ion optics system, discharge and neutralizer cathode assemblies, and the discharge chamber. Post-test inspection of the ion optics system revealed significant sputter erosion of the accelerator grid and measurable erosion of the screen grid. Post-test inspection results include the presence of through pits in the accelerator grid webbing, but no formation of rogue holes. Inspection of the discharge cathode indicates significant cathode orifice plate sputter erosion, as a result of exposure to the discharge plasma following removal of the keeper, but no erosion of or deposition near the orifice itself. Inspection of the neutralizer revealed no erosion to the keeper, orifice plate, or heater, and an orifice free of the deposits that were previously observed during the minimum power test segment. Post-test inspection indicates the discharge chamber experienced no measurable sputter erosion, although a substantial number of loose molybdenum coated carbon flakes were present on the sputter containment mesh. It is believed that the bulk of the flakes are due to back-sputtered beam target material, subsequently coated by sputter eroded grid material, a facility induced effect. A summary of the BOL and EOL performance, results of the system level inspection, an implication of findings to the ultimate service life capability of the 30-cm technology are presented.

MISSE 2 PEACE Polymers Atomic Oxygen Erosion Experiment on the International Space Station

Forty-one different polymer samples, collectively called the Polymer Erosion and Contamination Experiment (PEACE) Polymers, were exposed to the low Earth orbit (LEO) environment on the exterior of the International Space Station (ISS) for nearly 4 years as part of Materials International Space Station Experiment 2 (MISSE 2). The objective of the PEACE Polymers experiment was to determine the atomic oxygen erosion yield of a wide variety of polymeric materials after long-term exposure to the space environment. The polymers range from those commonly used for spacecraft applications, such as Teflon ® FEP, to more recently developed polymers, such as high temperature polyimide PMR (polymerization of monomer reactants). Additional polymers were included to explore erosion yield dependence upon chemical composition. The MISSE PEACE Polymers experiment was flown in MISSE Passive Experiment Carrier 2 (PEC 2), tray 1, attached to the exterior of the ISS Quest Airlock. It was exposed to atomic oxygen along with solar and charged particle radiation. MISSE 2 was successfully retrieved during a space walk on July 30, 2005 during Discoverys STS-114 Return to Flight mission. Details on the specific polymers flown, flight sample fabrication, pre-flight and post-flight characterization techniques, and atomic oxygen fluence calculations are discussed along with a summary of the atomic oxygen erosion yield results. The MISSE 2 PEACE Polymers experiment is unique because it has the widest variety of polymers flown in LEO for a long duration and was exposed to an unusually clean LEO spacecraft environment. This experiment provides extremely valuable erosion yield data for spacecraft design purposes.

Environmental Exposure Conditions for Teflon® Fluorinated Ethylene Propylene on the Hubble Space Telescope

The outer layer of Teflon® fluorinated ethylene propylene (FEP) multi-layer insulation (MLI) on the Hubble Space Telescope (HST) was observed to be significantly cracked at the time of the Second HST Servicing Mission (SM2), 6.8 years after HST was launched into low Earth orbit (LEO). Comparatively minor embrittlement and cracking were also observed in the FEP materials retrieved from solar-facing surfaces on the HST at the time of the First Servicing Mission (3.6 years exposure). After SM2, a failure review board was convened to address the problem of degradation of MLI on the HST. In order for this board to determine possible degradation mechanisms, it was necessary to consider all environmental constituents to which the FEP MLI surfaces were exposed. Based on measurements and various models, the environmental exposure conditions for the FEP surfaces on the HST were estimated, including: the number and temperature ranges of thermal cycles; equivalent sun hours; fluence and absorbed radiation dose of x-rays, trapped protons and electrons and plasma electrons and protons; and atomic oxygen (AO) fluence. This paper presents the environmental exposure conditions for FEP on the HST, briefly describing the possible roles of the environmental factors in the observed FEP embrittlement and providing references to the published works which describe in detail testing and analysis related to FEP degradation on the HST.

Hubble Space Telescope Metallized Teflon ® FEP thermal Control Materials: On-Orbit Degradation and Post-Retrieval Analysis

During the Hubble Space Telescope (HST) second servicing mission (SM2), degradation of unsupported Teflon® FEP (fluorinated ethylene propylene), used as the outer layer of the multilayer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon® FEP sample evaluation and additional testing of pristine Teflon® FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, ultraviolet and x-ray radiation along with thermal cycling. This paper discusses the testing and evaluation of the retrieved Teflon® FEP.

Insights into the Damage Mechanism of Teflon® FEP from the Hubble Space Telescope:

Metallized Teflon® FEP (fluorinated ethylene propylene) thermal control material on the Hubble Space Telescope (HST) has been found to be degrading in the space environment. Teflon® FEP thermal control blankets (space-facing FEP) retrieved during the first servicing mission (SM1) were found to be embrittled on solar-facing surfaces and contained microscopic cracks. During the second servicing mission (SM2) astronauts noticed that the FEP outer layer of the multi-layer insulation (MLI) covering the telescope was cracked in many locations around the telescope. Large cracks were observed on the light shield, forward shell and equipment bays. A tightly curled piece of cracked FEP from the light shield was retrieved during SM2 and was severely embrittled, as witnessed by ground testing. A failure review board was organized to determine the mechanism causing the MLI degradation. Density, x-ray crystallinity and solid-state nuclear magnetic resonance (NMR) analyses of the FEP retrieved during SM1 were inconsistent with results of FEP retrieved during SM2. Because the retrieved SM2 material was curled while in space, it experienced a higher temperature extreme during thermal cycling, estimated at 200°C, than the SM1 material, estimated at 50°C. An investigation on the effects of heating pristine FEP and FEP retrieved from the HST was therefore conducted. Samples of pristine, SM1 and SM2 FEP were heated to 200°C and evaluated for changes in density and morphology. Elevated-temperature exposure was found to have a major impact on the density of the retrieved materials. The characterization of the polymer morphology of the as-received and heated FEP by NMR provided results that were consistent with the density results. Differential scanning calorimetry (DSC) was conducted on pristine, SM1 and SM2 FEP. DSC results provided evidence of chain scission and increased crystallinity in the space exposed FEP, which supported the density and NMR results. Samples exposed to simulated solar flare x-rays, thermal cycling and long-term thermal exposure provided information on the environmental contributions to degradation. These findings have provided insight into the damage mechanisms of FEP in the space environment.

Atomic-Oxygen Undercutting of Protected Polymers in Low Earth Orbit

Hydrocarbon-based polymers that are exposed to atomic oxygen in low Earth orbit are slowly oxidized into volatile gases, which results in their erosion. Atomic-oxygen protective coatings that are both durable to atomic oxygen and effective in protecting underlying polymers have been developed. However, scratches, pin window defects, polymer surface roughness, and protective coating layer configuration can result in erosion and potential failure of protected thin polymer films even though the coatings are themselves atomic-oxygen durable. Issues are presented that cause protective coatings to become ineffective in some cases yet effective in others because of the details of their specific application. Observed in-space examples of failed and successfully protected materials using identical protective thin films are discussed and analyzed. Ground laboratory atomic-oxygen testing was conducted and compared with water vapor transport analyses from a previous study of protective coatings on Kapton® (polyimide), which indicates that vapor-deposited aluminized films are not as protective as sputter-deposited silicon dioxide films because of a greater number of pin window defects. Computational modeling was conducted and indicates that atomic-oxygen atoms trapped between the front and back surface aluminized films cause accelerated undercutting damage.

Effects of Radiation and Thermal Cycling on Teflon ® FEP:

Surfaces of the aluminized Teflon® FEP (fluorinated ethylene propylene) multilayer thermal insulation on the Hubble Space Telescope (HST) were found to be cracked and curled in some areas at the time of the second servicing mission (SM2) in February 1997, 6.8 years after HST was deployed in low Earth orbit (LEO). In an effort to understand what elements of the space environment might cause such damage, pristine second-surface aluminized Teflon® FEP was tested for durability to various types of radiation, to thermal cycling and to radiation followed by thermal cycling. Types of radiation included synchrotron vacuum ultraviolet and soft x-ray radiation, simulated solar flare x-ray radiation, electrons and protons. Thermal cycling was conducted in various temperature ranges to simulate HST orbital conditions for Teflon® FEP. Resultsoftensiletestingoftheexposedspecimensshowedthatexposuretohighfluencesofradiation caused degradation in tensile properties of FEP. However, exposure to radiation alone in exposures comparable to those experienced by HST did not produce reduction in ultimate tensile strength and elongation of Teflon® similar to that observed for HST-retrieved aluminized Teflon®. Synergism of radiation exposure and thermal cycling was evident in the results of three experiments: thermal cycling following electron and proton irradiation, thermal cycling following x-ray exposure, and additional thermal cycling of a sample retrieved from HST. However, irradiation and thermal cycling with comparable HST SM2 exposure conditions did not produce the degradation observed in the FEP material retrieved during HST SM2.

Prediction of In-Space Durability of Protected Polymers Based on Ground Laboratory Thermal Energy Atomic Oxygen

The probability of atomic oxygen reacting with polymeric materials is orders of magnitude lower at thermal energies (< 0.1 eV) than at orbital impact energies (4.5 eV). As a result, absolute atomic oxygen fluxes at thermal energies must be orders of magnitude higher than orbital energy fluxes, to produce the same effective fluxes (or same oxidation rates) for polymers. These differences can cause highly pessimistic durability predictions for protected polymers, and polymers which develop protective metal oxide surfaces as a result of oxidation if one does not make suitable calibrations. A comparison was conducted of undercut cavities below defect sites in protected polyimide Kapton samples flown on the Long Duration Exposure Facility (LDEF) with similar samples exposed in thermal energy oxygen plasma. The results of this comparison were used to quantify predicted material loss in space based on material loss in ground laboratory thermal energy plasma testing. A microindent hardness comparison of surface oxidation of a silicone flown on the Environmental Oxygen Interaction with Materials III (EOIM-III) experiment with samples exposed in thermal energy plasmas was similarly used to calibrate the rate of oxidation of silicone in space relative to samples in thermal energy plasmas exposed to polyimide Kapton effective fluences.

Mechanical Properties Degradation of Teflon ® FEP Returned From the Hubble Space Telescope

After 6.8 years on orbit, degradation has been observed in the mechanical properties of second-surface metalized Teflon® FEP (fluorinated ethylene propylene) used on the Hubble Space Telescope (HST) on the outer surface of the multi-layer insulation (MLI) blankets and on radiator surfaces. Cracking of FEP surfaces on HST was first observed upon close examination of samples with high solar exposure retrieved during the first servicing mission (SM1) conducted 3.6 years after HST was put into orbit. Astronaut observations and photographs from the second servicing mission (SM2), conducted after 6.8 years on orbit, revealed severe cracks in the FEP surfaces of the MLI on many locations around the telescope. This paper describes results of mechanical properties testing of FEP surfaces exposed for 3.6 years and 6.8 years to the space environment on HST. These tests include bend testing, tensile testing, and surface micro-hardness testing.

Deep Space 1 Flight Spare Ion Thruster 30,000-Hour Life Test

The extended-life test of the Deep Space 1 flight spare ion thruster was voluntarily terminated on 26 June 2003. During its five-year run, the thruster operated for a total of 30,352 h, processed 235.1 kg of xenon propellant, and demonstrated extended operation at multiple throttled conditions. The objectives of the test were to characterize failure modes and quantify thruster performance as a function of engine wear and throttle level. Degradation processes included erosion of the discharge cathode keeper, accelerator-grid sputter erosion, and deposition of material in the neutralizer cathode at low power. Performance degradation was limited to a reduction in measured thrust at the full-power point for the final 1000 h of operation. Posttest inspection of the enginewas initiated following the test termination to ascertain causes of the wear and to look for any previously unknown wear processes. Significant findings included facility-induced flakes in the discharge chamber, the presence of through-pits in the accelerator-grid webbing, significant erosion of the discharge cathode orifice plate, and healthy cathode inserts. A summary of the beginning-of-test and end-of-test performances and results of the posttest destructive evaluation are presented.