Engineering topics

Koontz, Steve

Publications and source records attributed to Koontz, Steve.

33 records · Page 2

NASA NTRS2010

Meeting Radiation Protection Requirements and Reducing Spacecraft Mass - A Multifunctional Materials Approach

Both crew and radio-sensitive systems, especially electronics must be protected from the effects of the space radiation environment. One method of mitigating this radiation exposure is to use passive-shielding materials. In previous vehicle designs such as the International Space Station (ISS), materials such as aluminum and polyethylene have been used as parasitic shielding to protect crew and electronics from exposure, but these designs add mass and decrease the amount of usable volume inside the vehicle. Thus, it is of interest to understand whether structural materials can also be designed to provide the radiation shielding capability needed for crew and electronics, while still providing weight savings and increased useable volume when compared against previous vehicle shielding designs. In this paper, we present calculations and analysis using the HZETRN (deterministic) and FLUKA (Monte Carlo) codes to investigate the radiation mitigation properties of these structural shielding materials, which includes graded-Z and composite materials. This work is also a follow-on to an earlier paper, that compared computational results for three radiation transport codes, HZETRN, HETC, and FLUKA, using the Feb. 1956 solar particle event (SPE) spectrum. In the following analysis, we consider the October 1989 Ground Level Enhanced (GLE) SPE as the input source term based on the Band function fitting method. Using HZETRN and FLUKA, parametric absorbed doses at the center of a hemispherical structure on the lunar surface are calculated for various thicknesses of graded-Z layups and an all-aluminum structure. HZETRN and FLUKA calculations are compared and are in reasonable (18% to 27%) agreement. Both codes are in agreement with respect to the predicted shielding material performance trends. The results from both HZETRN and FLUKA are analyzed and the radiation protection properties and potential weight savings of various materials and materials lay-ups are compared.

Atwell, William↗

NASA NTRS2009

Environmental Effects on ISS Materials Aging (1998 to 2008)

The performance of ISS spacecraft materials and systems on prolonged exposure to the low- Earth orbit (LEO) space flight are reported in this paper. In-flight data, flight crew observations, and the results of ground-based test and analysis directly supporting programmatic and operational decision-making are described. The space flight environments definitions (both natural and induced) used for ISS design, material selection, and verification testing are shown, in most cases, to be more severe than the actual flight environment accounting, in part, for the outstanding performance of ISS as a long mission duration spacecraft. No significant ISS material or system failures have been attributed to spacecraft-environments interactions. Nonetheless, ISS materials and systems performance data is contributing to our understanding of spacecraft material interactions with the spaceflight environment so as to reduce cost and risk for future spaceflight projects and programs. Orbital inclination (51.6 deg) and altitude (nominally near 360 km) determine the set of natural environment factors affecting the functional life of materials and systems on ISS. ISS operates in an electrically conducting environment (the F2 region of Earth s ionosphere) with well-defined fluxes of atomic oxygen, other charged and neutral ionospheric plasma species, solar UV, VUV, and x-ray radiation as well as galactic cosmic rays, trapped radiation, and solar cosmic rays. The LEO micrometeoroid and orbital debris environment is an especially important determinant of spacecraft design and operations. The magnitude of several environmental factors varies dramatically with latitude and longitude as ISS orbits the Earth. The high latitude orbital environment also exposes ISS to higher fluences of trapped energetic electrons, auroral electrons, solar cosmic rays, and galactic cosmic rays than would be the case in lower inclination orbits, largely as a result of the overall shape and magnitude of the geomagnetic field. As a result, ISS exposure to many environmental factors can vary dramatically along a particular orbital ground track, and from one ground track to the next, during any 24-hour period. The induced environment results from ISS interactions with the natural environment as well as environmental factors produced by ISS itself and visiting vehicles fleet. Examples include ram-wake effects, hypergolic thruster plume impingement, materials out-gassing, venting and dumping of fluids, and specific photovoltaic (PV) power system interactions with the ionospheric plasma (7-11). Vehicle size (L) and velocity (V), combined with the magnitude and direction of the geomagnetic field (B) produce operationally significant magnetic induction voltages (VxB.L) in ISS conducting structure during flight through high latitudes (> +45deg) during each orbit. Finally, an induced ionizing radiation environment is produced by cosmic ray interaction with the relatively thick ISS structure and shielding materials. The intent of this review article is, therefore, to provide a summary of selected aspects and elements of the ISS vehicle with regard to LEO space environment effects, associated with the much larger and more complicated vehicle that ISS has become since 1998, but also with an eye towards performance life extension to the year 2016 and beyond.

Alred, John↗

NASA NTRS2009

Using FLUKA to Calculate Spacecraft: Single Event Environments: A Practical Approach

The FLUKA nuclear transport and reaction code can be developed into a practical tool for calculation of spacecraft and planetary surface asset SEE and TID environments. Nuclear reactions and secondary particle shower effects can be estimated with acceptable accuracy both in-flight and in test. More detailed electronic device and/or spacecraft geometries than are reported here are possible using standard FLUKA geometry utilities. Spacecraft structure and shielding mass. Effects of high Z elements in microelectronic structure as reported previously. Median shielding mass in a generic slab or concentric sphere target geometry are at least approximately applicable to more complex spacecraft shapes. Need the spacecraft shielding mass distribution function applicable to the microelectronic system of interest. SEE environment effects can be calculated for a wide range of spacecraft and microelectronic materials with complete nuclear physics. Evaluate benefits of low Z shielding mass can be evaluated relative to aluminum. Evaluate effects of high Z elements as constituents of microelectronic devices. The principal limitation on the accuracy of the FLUKA based method reported here are found in the limited accuracy and incomplete character of affordable heavy ion test data. To support accurate rate estimates with any calculation method, the aspect ratio of the sensitive volume(s) and the dependence must be better characterized.

Koontz, Steve↗

NASA NTRS2007

Progress in Spacecraft Environment Interactions: International Space Station (ISS) Development and Operations

The set of spacecraft interactions with the space flight environment that have produced the largest impacts on the design, verification, and operation of the International Space Station (ISS) Program during the May 2000 to May 2007 time frame are the focus of this paper. In-flight data, flight crew observations, and the results of ground-based test and analysis directly supporting programmatic and operational decision-making are reported as are the analysis and simulation efforts that have led to new knowledge and capabilities supporting current and future space explorations programs. The specific spacecraft-environment interactions that have had the greatest impact on ISS Program activities during the first several years of flight are: 1) spacecraft charging, 2) micrometeoroids and orbital debris effects, 3) ionizing radiation (both total dose to materials and single event effects [SEE] on avionics), 4) hypergolic rocket engine plume impingement effects, 5) venting/dumping of liquids, 6) spacecraft contamination effects, 7) neutral atmosphere and atomic oxygen effects, 8) satellite drag effects, and 9) solar ultraviolet effects. Orbital inclination (51.6deg) and altitude (nominally between 350 km and 460 km) determine the set of natural environment factors affecting the performance and reliability of materials and systems on ISS. ISS operates in the F2 region of Earth s ionosphere in well-defined fluxes of atomic oxygen, other ionospheric plasma species, solar UV, VUV, and x-ray radiation as well as galactic cosmic rays, trapped radiation, and solar cosmic rays. The micrometeoroid and orbital debris environment is an important determinant of spacecraft design and operations in any orbital inclination. The induced environment results from ISS interactions with the natural environment as well as environmental factors produced by ISS itself and visiting vehicles. Examples include ram-wake effects, hypergolic thruster plume impingement, materials out-gassing, venting and dumping of fluids, and specific photovoltaic (PV) power system interactions with the ionospheric plasma. Vehicle size (L) and velocity (v), combined with the magnitude and direction of the geomagnetic field (B) produce operationally significant magnetic induction voltages (VxB.L) in ISS conducting structure during high latitude flight (>+/- 45deg) during each orbit. In addition, ISS is a large vehicle and produces a deep wake structure from which both ionospheric plasma and neutrals species are largely excluded. ISS must fly in a very limited number of approved flight attitudes, so that exposure of a particular material or system to environmental factors depends upon: 1) location on ISS, 2) ISS flight configuration, 3) ISS flight attitude, and 4) variation of solar exposure (Beta angle), and hence thermal environment, with time. Finally, an induced ionizing radiation environment is produced by trapped radiation and solar/cosmic ray interactions with the relatively massive ISS structural shielding.

Koontz, Steve↗

NASA NTRS2006

Impact of Solar Array Position on ISS Vehicle Charging

The International Space Station (ISS), because of its large structure and high voltage solar arrays, has a complex plasma interaction with the ionosphere in low Earth orbit (LEO). This interaction of the ISS US Segment photovoltaic (PV) power system with the LEO ionospheric plasma produces floating potentials on conducting elements of the ISS structure relative to the local plasma environment. To control the ISS floating potentials, two Plasma Contactor Units (PCUs) are installed on the Z1 truss. Each PCU discharges accumulated electrons from the Space Station structure, thus reducing the potential difference between the ISS structure and the surrounding charged plasma environment. Operations of the PCUs were intended to keep the ISS floating potential to 40 Volts (Reference 1). Exposed dielectric surfaces overlying conducting structure on the Space Station will collect an opposite charge from the ionosphere as the ISS charges. In theory, when an Extravehicular Activity (EVA) crewmember is tethered to structure via the crew safety tether or when metallic surfaces of the Extravehicular Mobility Unit (EMU) come in contact with conducting metallic surfaces of the ISS, the EMU conducting components, including the perspiration-soaked crewmember inside, can become charged to the Space Station floating potential. The concern is the potential dielectric breakdown of anodized aluminum surfaces on the EMU producing an arc from the EMU to the ambient plasma, or nearby ISS structure. If the EMU arcs, an electrical current of an unknown magnitude and duration may conduct through the EVA crewmember, producing an unacceptable condition. This electrical current may be sufficient to startle or fatally shock the EVA crewmember (Reference 2). Hence, as currently defined by the EVA community, the ISS floating potential for all nominal and contingency EVA worksites and translation paths must have a magnitude less than 40 volts relative to the local ionosphere at all times during EVA. Arcing from the EMU is classified as a catastrophic hazard, which requires two-failure tolerant controls, i.e., three hazard controls. Each PCU is capable of maintaining the ISS floating potential below the requirement during EVA. The two PCUs provide a single failure tolerant control of ISS floating potential. In the event of the failure of one or two PCUs, a combination of solar array shunting and turning the solar arrays into their own wakes will be used to supply control of the plasma hazard (Reference 3). The purpose of this paper is to present on-orbit information that shows that ISS solar array placement with respect to the ISS velocity vector can control solar array plasma charging, and hence, provide an operational control for the plasma hazard. Also, this paper will present on-orbit information that shows that shunting of the ISS solar arrays can control solar array plasma charging, and hence, provide an additional operational control for the plasma hazard.

Alred, John↗

NASA NTRS2006

Analysis of ISS Plasma Interaction

To date, the International Space Station (ISS) has been one of the largest objects flown in lower earth orbit (LEO). The ISS utilizes high voltage solar arrays (160V) that are negatively grounded leading to pressurized elements that can float negatively with respect to the plasma. Because laboratory measurements indicate a dielectric breakdown potential difference of 80V, arcing could occur on the ISS structure. To overcome the possibility of arcing and clamp the potential of the structure, two Plasma Contactor Units (PCUs) were designed, built, and flown. Also a limited amount of measurements of the floating potential for the present ISS configuration were made by a Floating Potential Probe (FPP), indicating a minimum potential of 24 Volts at the measurement location. A predictive tool, the ISS Plasma Interaction Model (PIM) has been developed accounting for the solar array electron collection, solar array mast wire and effective conductive area on the structure. The model has been used for predictions of the present ISS configuration. The conductive area has been inferred based on available floating potential measurements. Analysis of FPP and PCU data indicated distribution of the conductive area along the Russian segment of the ISS structure. A significant input to PIM is the plasma environment. The International Reference Ionosphere (IRI 2001) was initially used to obtain plasma temperature and density values. However, IRI provides mean parameters, leading to difficulties in interpretation of on-orbit data, especially at eclipse exit where maximum charging can occur. This limits our predicative capability. Satellite and Incoherent Scatter Radar (ISR) data of plasma parameters have also been collected. Approximately 130,000 electron temperature (Te) and density (Ne) pairs for typical ISS eclipse exit conditions have been extracted from the reduced Langmuir probe data flown aboard the NASA DE-2 satellite. Additionally, another 18,000 Te and Ne pairs of ISR data from several radar locations around the globe were used to assure consistency of the satellite data. PIM predictions for ISS charging made with this data correlated very well with FPP data, indicating that the general physics of spacecraft charging with high voltage solar arrays have been captured. The predictions also provided the probabilities of occurrences for ISS charging. These probabilities give a numerical measure of the number of times when the ISS will approach or exceed the vehicle plasma hazard conditions for each configuration. In this paper we shall present the interaction mechanisms between the ISS and the surrounding plasma and give an overview of the PIM components. PIM predictions are compared with available data followed by a discussion of the variability of plasma parameters and the conductive area on the ISS. The ISS PIM will be further tested and verified as data from the Floating Potential Measurement Unit become available, and construction of the ISS continues.

Reddell, Brandon↗

NASA NTRS2005

Simulated Space Environment Effects on the Blocking Force of Silicone Adhesive

The International Space Station (ISS) solar arrays utilize MD-944 diode tape to protect the underlying diodes in the solar array panel circuit and also provide thermal conditioning and mechanical support. The diode tape consists of silicone pressure sensitive adhesive (Dow Coming QC-7725) with a protective Kapton over-layer. On-orbit, the Kapton over-layer will erode under exposure to atomic oxygen (AO) and the underlying exposed silicone adhesive will ultimately convert, under additional AO exposure, to a glass like silicate. The current operational plan is to retract ISS solar array P6 and leave it stored under load for a long duration (6 months or more) during ISS assembly. With the Kapton over-layer eroded away, the exposed silicone adhesive must not cause the solar array to stick to itself or cause the solar array to fail during redeployment. Previous testing by Lockheed-Martin Space Systems (LMSS) characterized silicone blocking following exposure to low energy atomic oxygen (AO) in an asher facility, but this is believed to be conservative. An additional series of tests was performed by the Environmental Effects Group at MSFC under direction from the ISS Program Office Environments Team. This test series included high energy AO (5 eV), near ultraviolet (NUV) radiation and ionizing radiation, singly and in combination. Additional samples were exposed to thermal energy AO (<0.1 ev) for comparison to the LMSS tests. Diode tape samples were exposed to each environment constituent individually, put under preload for seven days and then the resulting blocking force was measured using a tensile machine. Additional samples were exposed to AO, NUV and electrons in series and then put under long term (three to ten months) preload to determine the effect of preload duration on the resulting blocking force of the silicone-to-silicone bond. Test results indicate that high energy AO, ultraviolet radiation and electron ionizing radiation exposure all reduce the blocking force for a silicone-to-silicone bond. AO exposure produces the most significant reduction in blocking force.

Boeder, Paul↗

NASA NTRS2004

Fuel Oxidizer Reaction Products (FORP) Contamination of Service Module (SM) and Release of N-nitrosodimethylamine(NDMA)in a Humid Environment from Crew EVA Suits Contaminated with FORP

The Service Module (SM) is an element of the Russian Segment of the International Space Station (ISS). One of the functions of the SM is to provide attitude control for the ISS using thrusters when the U.S. Control Moment Gyros (CMG's) must be desaturated. Prior to an Extravehicular Activity (EVA) on the Russian Segment, the Docking Compartment (DC1) is depressurized, as it is used as an airlock. When the DC1 is depressurized, the CMG's margin of momentum is insufficient and the SM attitude control thrusters need to fire to desaturate the CMG's. SM roll thruster firings induce contamination onto adjacent surfaces with Fuel Oxidizer Reaction Products (FORP). FORP is composed of both volatile and non-volatile components. One of the components of FORP is the potent carcinogen N-nitrosdimethylamine (NDMA). Since the EVA crewmembers often enter the area surrounding the thrusters for tasks on the aft end of the SM and when translating to other areas of the Russian Segment, the presence of FORP is a concern. This paper will discuss FORP contamination of the SM surfaces, the release of NDMA in a humid environment from crew EVA suits, if they happen to be contaminated with FORP, and the toxicological risk associated with the NDMA release.

Schmidl, William↗

NASA NTRS2004

Assessment and Control of Spacecraft Charging Risks on the International Space Station

The International Space Station (ISS) operates in the F2 region of Earth's ionosphere, orbiting at altitudes ranging from 350 to 450 km at an inclination of 51.6 degrees. The relatively dense, cool F2 ionospheric plasma suppresses surface charging processes much of the time, and the flux of relativistic electrons is low enough to preclude deep dielectric charging processes. The most important spacecraft charging processes in the ISS orbital environment are: 1) ISS electrical power system interactions with the F2 plasma, 2) magnetic induction processes resulting from flight through the geomagnetic field and, 3) charging processes that result from interaction with auroral electrons at high latitude. Recently, the continuing review and evaluation of putative ISS charging hazards required by the ISS Program Office revealed that ISS charging could produce an electrical shock hazard to the ISS crew during extravehicular activity (EVA). ISS charging risks are being evaluated in an ongoing measurement and analysis campaign. The results of ISS charging measurements are combined with a recently developed model of ISS charging (the Plasma Interaction Model) and an exhaustive analysis of historical ionospheric variability data (ISS Ionospheric Specification) to evaluate ISS charging risks using Probabilistic Risk Assessment (PRA) methods. The PRA combines estimates of the frequency of occurrence and severity of the charging hazards with estimates of the reliability of various hazard controls systems, as required by NASA s safety and risk management programs, to enable design and selection of a hazard control approach that minimizes overall programmatic and personnel risk. The PRA provides a quantitative methodology for incorporating the results of the ISS charging measurement and analysis campaigns into the necessary hazard reports, EVA procedures, and ISS flight rules required for operating ISS in a safe and productive manner.

Koontz, Steve↗

NASA NTRS2003

Assessment and Control of Spacecraft Charging Risks on the International Space Station

Electrical interactions between the F2 region ionospheric plasma and the 160V photovoltaic (PV) electrical power system on the International Space Station (ISS) can produce floating potentials (FP) on the ISS conducting structure of greater magnitude than are usually observed on spacecraft in low-Earth orbit. Flight through the geomagnetic field also causes magnetic induction charging of ISS conducting structure. Charging processes resulting from interaction of ISS with auroral electrons may also contribute to charging albeit rarely. The magnitude and frequency of occurrence of possibly hazardous charging events depends on the ISS assembly stage (six more 160V PV arrays will be added to ISS), ISS flight configuration, ISS position (latitude and longitude), and the natural variability in the ionospheric flight environment. At present, ISS is equipped with two plasma contactors designed to control ISS FP to within 40 volts of the ambient F2 plasma. The negative-polarity grounding scheme utilized in the ISS 160V power system leads, naturally, to negative values of ISS FP. A negative ISS structural FP leads to application of electrostatic fields across the dielectrics that separate conducting structure from the ambient F2 plasma, thereby enabling dielectric breakdown and arcing. Degradation of some thermal control coatings and noise in electrical systems can result. Continued review and evaluation of the putative charging hazards, as required by the ISS Program Office, revealed that ISS charging could produce a risk of electric shock to the ISS crew during extra vehicular activity. ISS charging risks are being evaluated in ongoing ISS charging measurements and analysis campaigns. The results of ISS charging measurements are combined with a recently developed detailed model of the ISS charging process and an extensive analysis of historical ionospheric variability data, to assess ISS charging risks using Probabilistic Risk Assessment (PRA) methods. The PRA analysis (estimated frequency of occurrence and severity of the charging hazards) are then used to select the hazard control strategy that provides the best overall safety and mission success environment for ISS and the ISS crew. This paper presents: 1) a summary of ISS spacecraft charging analysis, measurements, observations made to date, 2) plans for future ISS spacecraft charging measurement campaigns, and 3) a detailed discussion of the PRA strategy used to assess ISS spacecraft charging risks and select charging hazard control strategies

Koontz, Steve↗

NASA NTRS1993

Shuttle primary reaction control system engine exhaust plume contamination effects

Space Shuttle proximity operations constitute an important part of the SSF induced external environment. The impingement of primary reaction control system (PRCS) engine plumes on SSF functional surfaces during docking or berthing and separation leads to concerns about molecular contamination and high speed particle impact. The Shuttle Plume Impingement flight Experiment (SPIE) was designed to provide a direct measure of both the molecular contamination and particle impact rates produced by Shuttle PRCS engines in the LEO environment. The measured permanent deposition produced by PRCS engine firings was less than that assumed in current SSF programatic assessments. Only two to three possible high velocity particle impact pits were observed on the RMS end effector hardware.

Koontz, Steve↗

NASA NTRS1992

Recommended practices for in-space and ground laboratory. Atomic oxygen exposure and analysis

A detailed guide to testing materials for atomic oxygen durability in earth orbit environments is presented. The steps covered include sample preparation, including masking of the sample, dehydration, weighing, and handling; effective fluence prediction, including the use of witness samples (notably Kapton); plasma facility and operational considerations, involving such matters as avoidance of silicone contamination, the use of continuous versus incremental ashing, and temperature of operation; and erosion yield measurement, with calculation methods and protective coating performance indices provided.

Banks, Bruce↗

NASA NTRS1991

Chemically Layered Porous Solids

Aerogels and other porous solids in which surfaces of pores have chemical properties varying with depth below macroscopic surfaces prepared by sequences of chemical treatments. Porous glass or silica bead treated to make two depth zones having different chemical properties. Beads dropped along tube filled with flowing gas containing atomic oxygen, generated in microwave discharge. General class of materials treatable include oxides of aluminum, silicon, zirconium, tin, titanium, and nickel, and mixtures of these oxides. Potential uses of treated materials include chromatographic separations, membrane separations, controlled releases of chemicals, and catalysis.

Koontz, Steve↗

NASA NTRS1991

Monatomic-Oxygen Reactors For Materials Testing And Surface Chemistry

Two kinds of gas-flow reaction chambers expose specimens of materials to spatially even fluxes of monatomic oxygen. Monatomic oxygen diffuses along side-arm tubes attached to flowing-after-glow apparatus. If rate of diffusion high enough, specimens at ends of these tubes receive uniform doses. Monatomic oxygen generated by photolysis in pulsed laser beam diffuses sideways out of laser beam and onto nearby surfaces of specimens.

Koontz, Steve↗

NASA NTRS1988

IR fiber temperature sensing system

Infrared fiber optic pyrometry has become a practical reality using improved strength fluoride glass fibers. The addition of a plastic coating and rugged cabling allows the fibers to be used in the field. A detailed theoretical model of the infrared fiber optic pyrometer (non-contact or radiative thermometer) has been derived and compared with data produced by a prototype fluoride glass fiber radiative thermometer. Excellent agreement was obtained between theory and experiment over a temperature range of 30 to 700 C.

Tran, D. C.↗