International Conference on Environmental Systems

Permanent URI for this collectionhttps://hdl.handle.net/2346/58495

The International Conference on Environmental Systems, or ICES (known prior to 1990 as the Intersociety Conference on Environmental Systems), is an annual technical conference focusing on human spaceflight technology and space human factors. Session topics include: Environmental Control and Life Support Systems (ECLSS), thermal control, life sciences, extra-vehicular activity (EVA) systems (including space suit design and human-robot interaction), space architecture, and mission planning for exploration. The conference has taken place annually since 1971.

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Comparison of Artemis I Radiation Measurements with Orion EFT-1 and ISS Data
(2024 International Conference on Environmnetal Systems, 2024-07-21) Gaza, Ramona; Stoffle, Nicholas; Campbell-Ricketts, Thomas; George, Stuart; Semones, Edward; Dimapilis, Dinah
Comparison of Artemis I Radiation Measurements with Orion EFT-1 and ISS Data Ramona Gaza, Ph.D. On behalf of the Space Radiation Analysis Group Leidos, Space Exploration and Mission Operations, Houston, TX 77058, USA Space Radiation Analysis Group, NASA Johnson Space Center, Houston, TX 77058, USA Corresponding author: ramona.gaza-1@nasa.gov The first major spaceflight of NASA's Artemis program to return humans to the Moon, the Artemis I uncrewed mission, has been flown successfully November 16 - December 11, 2022, for a total mission duration of 25.5 days. The Space Radiation Analysis Group (SRAG) at NASA Johnson Space Center (JSC) has provided a suit of passive radiation detectors and active instruments in support of multiple Artemis I Science Payloads. The same passive technology has been flying in support of the International Space Station (ISS) for more than 20 years and has been successfully flown on the NASA Orion Exploration Flight Test 1 (EFT-1) launched on December 5, 2014, with a duration of only 4.5 hours. The Orion EFT-1 trajectory included two orbits around the Moon with a high apogee which was different from the from the Artemis I distant retrograde orbit trajectory, resulting in a significant radiation exposure difference through the Van Allen belts. On ISS, the average daily dose is modulated by the 11-year solar cycle and solar minimum dose values will differ from solar maximum daily doses. This presentation will include an overview of the Artemis 1 science payloads radiation data in comparison with the Orion EFT-1 and ISS measurements.
Update on Active Shielding Concept Using Electrostatic Fields
(2024 International Conference on Environmnetal Systems, 2024-07-21) Madzunkov, Stojan; Nikolic, Dragan; Fry, Dan; Bahadori, Amir; Chowdhury, Rajarshi Pal; Stegeman, Luke; Arnett, Kenneth; Battel, Steven; Hancock, Allison; Gilchrist, Brian; Leon, Omar; McNally, Patrick; Lund, Matthew; Delzanno, Gian Luca
We have investigated different electrostatic field configurations to identify the optimal configuration to maximize the number of safe days astronauts can be exposed to ionizing radiation (SEP and CGR) during exploration missions. The general concept is based on principles of charged particle shielding by Earth's geomagnetic field, i.e., particle deflection. We have utilized custom-built simulation tools for fast GPU-based computing of different configurations and measurements performed at the Brookhaven National Lab Tandem facility using an analogous "wind tunnel" setup. This has allowed us to develop scaling laws by directly measuring the shielding efficacy of scaled-down three-dimensional test articles, resulting in methods to mature technology to effectively scale up to both in-space particle energies and physical shield size (mass, power) required to reduce the cumulative radiation exposure to humans. In addition, work was conducted to develop ways to mitigate the interaction with the in-space plasma environment. Measurements were performed on several mitigation methods. Results showed the possibility of reducing power requirements from megawatts to tens of kilowatts. Lastly, a prototype high-voltage power supply design capable of reaching 300 kV was investigated and verified with a simple single electrostatic dipole configuration (SPRL). We find that the currently identified configuration will reduce the dose from the 1989 SPE by approximately 50% with an applied voltage of 1 MV and a power consumption of ~10 kW. The next step in concept maturation is a planned large-scale trade study to identify required support structures, electrode charging scenarios, operational concepts, etc., to implement and operate the concept as a full-size three-dimensional shield around a Mars transit vehicle or surface habitat. This trade will allow for mass/power estimates of a full-scale shield and the identification of current technologies to support the construction of an active radiation shield.
MELiSSA Pilot Plant Regenerative Life Support System
(2024 International Conference on Environmnetal Systems, 2024-07-21) Godia, Francesc; Arnau, Carolina; Ciurans, Carles; Vilaplana, Marcel; Peiro, Enrique; Dusssap, Claude-Gilles; Poughon, Laurent; Gerbi, Olivier; Pannico, Antonio; de Pascale, Stefania; Lamaze, Brigitte; Audas, Chloé; Lasseur, Christophe; Vizcarra, Arnau
MELiSSA is developing regenerative Life Support technologies for long-term Space missions. The MELiSSA loop is conceived as several compartments, each one performing a specific function, providing all together the basic functions in life support: food production, atmosphere regeneration, water recovery and waste treatment. The MELiSSA Pilot Plant is a facility designed for the terrestrial demonstration of this concept, hosting laboratory rats as a crew mimicking the respiration of humans, working at terrestrial conditions in industry standards and long-term experiments. Today, the work at the Pilot Plant is focused on the connection of compartments in continuous and controlled operation, in several consecutive steps, targeting to the completion of the loop. Current activities are focused on the integration of four compartments: Compartment 3 (nitrifying packed-bed bioreactor based on the co-culture of immobilized Nitrosomonas winogradsky and Nitrobacter europaea), Compartment 4a (an air-lift photobioreactor for the culture of the edible cyanobacteria Limnospira indica with concomitant oxygen production), Compartment 4b (a 5 m2 hydroponic culture plant chamber working with lettuce as plant model, producing O2 and edible material) and Compartment 5 (an animal isolator with rats as mock-up crew). The output from several continuous operation and long-term integration experiments under controlled conditions will be reported. The system showed high robustness and reliability over long operation periods (several months). The performance of the system has been analyzed both in steady-state and dynamics conditions, with extensive on-line instrumentation for the main variables, under different experimental. The transition of the system to the use of human urine is also in progress. Overall, the MELiSSA Pilot Plant has consolidated its program over the last years and is producing a collection of data demonstrating the capabilities of MELiSSA as a regenerative life support system. Future steps in the evolution of this facility will also be presented.
Orion LAMS Laser Absorption Spectrometer for Human Spaceflight – Artemis 4 and 5 Design Updates and Flight Builds
(2024 International Conference on Environmnetal Systems, 2024-07-21) Pohly, Jason; Roe, David; Erb, Cody; Mansour, Kamjou; Christensen, Lance
The Orion Laser Air Monitor System (LAMS) is a tunable laser spectrometer that will monitor oxygen, carbon dioxide, and water vapor levels in the Orion Multipurpose Crew Vehicle (MCPV) cabin and in the space suit loop. LAMS, designed to be small, lightweight, and low power, can nonetheless accurately measure a wide dynamic range of analyte concentrations over relatively wide pressure and temperature ranges despite not using gas pumps, flow, or pressure controllers. Additionally, the LAMS hardware and electronics are capable of meeting stringent Crit-1R requirements for human life support. This paper is a follow-up to the 2020 and 2023 ICES papers which covered flight unit build and testing results for Artemis missions 2 and 3. This paper covers design updates, flight unit build, and test results for the Artemis missions 4 and 5.
Qualification of Thermal Hardware for the Europa Clipper Mission
(2024 International Conference on Environmnetal Systems, 2024-07-21) Schmidt, Tyler; Hunter, Madison
The Europa Clipper mission and environment present multiple challenges for design and implementation of flight hardware. The biggest environmental driver is ionizing radiation around Jupiter. The expected radiation total ionizing dose on external surfaces is over 1 Grad, and there are related internal electrostatic discharge effects as well. The mission trajectory solar distance varies between 0.82 and 5.6 AU, which can create large temperature variations over the mission lifetime. Some of the science instruments impose contamination control and magnetic cleanliness requirements. Lastly, the demands of this flagship mission require high performance and low risk hardware for mission success. Due to the confluence of these requirements and the demanding needs for many of them, all thermal hardware required appropriate qualification analysis and testing even if successfully flown on a previous mission. Flight heritage hardware often served as a starting point for design and implementation approaches. Those designs were then adapted to suit Europa Clipper. Qualification testing for radiation, internal electrostatic discharge, magnetic emissions, and thermal cycling served as important confirmations of acceptability (or not) of hardware for flight. Other evaluations were sometimes performed for material analysis, contamination control, or parts reliability. This paper will summarize qualification analysis and testing motivation, parameters, results, and lessons learned for various thermal hardware that will fly on Europa Clipper.
Thermal Control Design of the ESCAPADE Mars Space Vehicles
(2024 International Conference on Environmnetal Systems, 2024-07-21) Reilly, Sean; Arevalo-Londona, Daniel; Huang, Steven; Vincent, Taylor; Scherf, Karl; Muy, Terry; Tarantini, Vincent; Cepeda-Rizo, Juan
The Escape and Plasma Acceleration and Dynamics Explorers (EscaPADE) are a pair of spacecraft manufactured by Rocket Lab USA's Space Systems divisions, overseen by the Space Sciences Lab at UC Berkeley. The primary science instruments are intended to study the interaction between ions and Mars' magneto-sphere. The identical space vehicles experience a wide variety of thermal environments at Earth, on the way to Mars, and at Mars. Additionally, the spacecraft are expected to experience solar conjunction during cruise, further complicating the fault protection scheme. This work will detail the thermal challenges related to a pair relatively low mass Mars orbiters performing apogee raising maneuvers at Earth, orbital capture at Mars, and multiple burns in orbit around Mars over the prime science mission of the orbiters. The space vehicles will have to keep the bi-propellant propulsion system active and ready to change orbital parameters at roughly 6 month intervals over the 3 year prime science mission presenting significant thermal challenges. The Authors will discuss some of the problems that had to be overcome in the design as well as some innovative solutions to minimize complexity and maximize resources to deliver a robust Mars bound thermal design.
Overview of Integrated Random Vibration Testing of the NASA Orion Crew Survival Suit
(2024 International Conference on Environmnetal Systems, 2024-07-21) Suhey, Jeffrey; Gohmert, Dustin; Baldwin, Mark
The launch ascent and abort random vibration environments from NASA's Orion spacecraft drove the need to test the NASA Orion Crew Survival Suit as an integrated system with exposure to the design levels. In order to properly characterize component responses, a series of integrated tests were designed to incorporate the interaction between a crew member, suit, seat, and the attenuation system. The first Development Test (2017) included human subjects, and was performed with early development seat and suit components, and low level inputs. It provided a baseline for frequency response and behavior of the integrated system. The next two tests, Development Test (2019) and Qualification Test (2020), used manikin surrogates to represent the crew member and increasing levels of component hardware fidelity in order to test to higher input levels. Testing was performed at NASA Johnson Space Center (JSC) and Kennedy Space Center (KSC) Vibration Labs and provided the input levels required to represent the ascent and abort vehicle profiles as well as recorded component response behavior from accelerometer instrumentation and high speed cameras. Inspection of the suit and related components showed that for all seat orientations, and input environments, no damage occurred. Additional pre and post-test checks confirmed the functionality of all suited hardware. Response data of the suited components generally showed heavy attenuation across most of the tested frequency range. Transmissibility plots showed some amplification of components at lower frequency ranges. Overall this series of integrated tests showed that 1) the use of surrogate manikins in the tests were adequate for representing crew in a vibration environment, 2) full vibration levels for ascent and abort were heavily attenuated in the suit components and were non-damaging, and 3) the suit and related components are qualified for the Orion random vibration environments.
Advanced Pressure Garment Space Suit Sizing Considerations
(2024 International Conference on Environmnetal Systems, 2024-07-21) Jerome, Christine; Rhodes, Richard; Campbell, Donald
Fitting a space suit to a person could be considered an art form. Establishing a repeatable fitcheck process to accommodate the full anthropometric range of test subjects for a single suit design is a critical process to be able to prove suit functionality. Suited test subjects can have different preferences on how they fit inside a suit, and different suited test environments can lead to differences in certain suit sizing accommodations. Throughout this paper, the process of achieving an acceptable suit fit for test subjects will be discussed, along with sizing considerations for changes in suit design and test environments. Lessons learned from the Exploration Extravehicular Mobility Unit (xEMU) will be used to describe specific examples, along with key takeaways from additional NASA prototype mobility space suits. The suit fitcheck process starts with utilizing the anthropometric measurements of a test subject to evaluate their relation to the dimensions of space suit hardware in creating a predicted suit sizing configuration. From here, subjective comments and test team observations drive iterations to the suit sizing configuration to culminate in an acceptable suit fit for performance of further test evaluations. Understanding the relation of subjective comments to their impact on altering suit sizing is critical in establishing an acceptable suit fit.
Durable Anti-Reflective and Anti-Fog Coatings Produced by Aerosol Impact Driven Assembly
(2024 International Conference on Environmnetal Systems, 2024-07-21) Firth, Peter; Holman, Zachary; Matthews, Dave; Newhouse-Illige, Ty; Ramirez, Ronnie; Victoria, Albert
The Exploration Extravehicular Mobility Unit's (xEMU) helmet is a complex assembly designed to accomplish several critical tasks. In addition to maintaining a suitable environment for the wearer, it must also allow for appropriate mobility and provide a wide and undistorted view of the surroundings. A critical component of the helmet's optical system is the anti-fog coating. While previous versions of the anti-fog coating have provided suitable anti-fog performance, they have been difficult to apply, lacked mechanical/chemical durability, and/or resulted in unanticipated failures (e.g., outgassing of eye-irritating materials during use). This work will describe the use of a new coating technology, aerosol impact driven assembly (AIDA), to develop the next generation permanent anti-fog coating for the xEMU helmet. We will use AIDA's unique ability to tune both the refractive index and surface roughness of the film to deposit a thin (<150 nm), transparent (>85% transmittance of visible light), and hydrophilic (contact angle <10°) anti-fog coating. We will characterize the anti-fog, optical, and durability (abrasion and chemical resistance) performance on polycarbonate substrates up to 15" x 20" -- the size required to manufacture operational helmets. We will conclude with an analysis on the path towards commercial production of the coating and discuss other potential use-cases within and beyond the xEMU helmet.
Thermal Design for the New Era of Lunar Instruments: Lessons Learned from NASA Goddard Space Flight Center’s Instrument Design Laboratory
(2024 International Conference on Environmnetal Systems, 2024-07-21) Yang, Kan; Matson, Elizabeth; Peabody, Hume; Choi, Michael; Steinfeld, David; Ottens, Brian; Coronado, Patrick; Appelbaum, Scott
As NASA enters the Age of Artemis, the resurgence of interest in the Moon has not only contributed to rapid development of crewed vehicles and habitats, it has also inspired a design renaissance for robotic and astronaut-deployed instrumentation for the lunar surface. NASA Goddard Space Flight Center's Instrument Design Laboratory (IDL) has been at the forefront of lunar instrument concept design, and in recent years has used its collaborative and concurrent process to design instruments ranging from small devices held or worn by astronauts to large lander- or rover-mounted instruments with multiple scientific or technical functions. This paper explores the wide spectrum of lunar instruments studied by the IDL within the past few years and their thermal design nuances. A series of lessons learned will be used to categorize the general trends prevalent across lunar-specific design and the thermal and systems-level drivers for instrumentation on the surface of the Moon. While the instruments investigated in this work are not named due to their competition sensitivity and proprietary nature, it is hoped that the lessons learned captured here will help identify specific design drivers and areas of focus for the thermal and systems engineers designing the next generation of lunar instrumentation.
Wire-break Ignition Testing of Materials for Spacesuit Fire Hazard Control
(2024 International Conference on Environmnetal Systems, 2024-07-21) Campbell, Colin; Peralta, Stephen; Ward, Virginia; Abney, Morgan; Morris, Danielle; Gallus, Tim
The input design constraints applied to the Extra-Vehicular Activity (EVA) spacesuit pose a significant challenge for mitigation of the fire hazard. In order to minimize fatigue and increase comfort of the crewmember operating the suit, the suit pressure is lowered below sea level conditions with typical EVA suit designs operating with 4.3 psia (29.7 kPa). With the lowered operating pressures and the use of closed loop life support, the suit requires elevated concentrations of oxygen typically >95%. At these oxygen concentrations, nearly the entirety of the suit internal materials are flammable. This leaves one remaining possible control leg of the fire triangle: ignition sources. Since the Gemini and Apollo programs, this has been a risk that has under constant reassessment with focus on improved mitigation. After the Apollo I fire, an arc ignition method was developed and used to quantify ignition thresholds for in-suit materials resulting in a current limit for powered in-suit devices applied to all suit designs that followed. After the discovery of a frayed spacesuit biomed cable on STS-113 during the Shuttle Program, the previous arcing method was repeated with additional methods developed to extend the testing further. One of those methods was Wire-break Ignition Testing in which the current in a single strand of wire was progressively taken higher preheating the material in proximity with a resultant break and the application of a reasonably repeatable arc to ignite the material all while exposed in the selected environment. This method was used to test a suite of spacesuit materials providing relative performance with respect to ignition with this particular configuration for application of energy. Leveraging the previous data and extending it further to consider more and recent materials coupled with lowering energy levels via smaller wire gauges is the subject of this paper.
Electrostatic Discharge Hazard in Spacesuits
(2024 International Conference on Environmnetal Systems, 2024-07-21) Campbell, Colin; Buhler, Charles; Abney, Morgan; Toth, Joseph
Triboelectric charging with a resultant electrostatic discharge is a phenomenon with which most people are familiar, especially those living in dry climates. Spacesuits, which are elaborately designed anthropomorphic pressure vessels attempting to match the motions of the human body while providing protections from the harsh environment of space, also must contend with triboelectric charging. But for spacesuits, which operate with elevated oxygen concentrations (usually >95%) in order to enable the lowest reasonable working pressure for the human operator, additional challenges are present with potentially catastrophic results if those challenges are not met successfully. A painful spark discharge can offer as much as 15kV with an energy transfer as high as 15-20mJ. This is well above that needed to ignite hydrocarbon vapors or fine dusts such as those in grain silos. There have been several test methods beyond the scope of this paper seeking to address the mechanistic ignition energies to ignite suit materials with a wide possible range of results due to the varied test configurations. What is not known is if the suit materials offer sufficient capability for triboelectric charge generation and discharge to generate the potential and the resultant discharge energies within the ranges needed to ignite the suit materials in this operating environment.
The Effect of Ambient Pressure and External Heating on the Limiting Oxygen Concentration (LOC) of PMMA Rods Caused by Oxygen Depletion in Confined Fires
(2024 International Conference on Environmnetal Systems, 2024-07-21) Fernandez-Pello, Carlos; Etzenbach, Lilly; Liveretou, Christina; Rivera, Jose; Gollner, Michael; Olson, Sandra; Ferkul, Paul
Studying the impact of different ambient environmental conditions is important to characterize the different processes involved in solid combustion and to dictate fire safety standards adapted to extreme environments, such as Space Exploration Atmospheres (SEA). The environments of spacecraft and space facilities are characterized by reduced gravity, low flow velocities, low pressures, and elevated oxygen concentrations. This work aims to assess the flammability limits of solid combustibles burning in confined environments. Specifically, we study the Limiting Oxygen Concentration (LOC) of flames spreading downward over polymethyl methacrylate (PMMA) cylinders in an opposed gas flow of decreasing oxygen concentration. This is done by analyzing the natural depletion of oxygen due to the fuel burning under subatmospheric pressures, external radiation, and mixed flow conditions. Results show that the LOC increases as the pressure decreases and decreases as the external heat fluxes increase. Adding a source of opposed forced flow subsequently increases the LOC, although the effect becomes more negligible at higher pressures and higher radiant fluxes, as buoyancy becomes the prominent flow driver. These results are interpreted through a study of chemical kinetics mechanisms, focusing on the Damkholer number. The data of this study is relevant in the determination of the flammability of combustible materials in confined spaces such as airplanes, submarines, and spacecraft. As per the latter, the data will later be compared to that collected through experiments on the International Space Station (ISS) under the SoFIE-MIST project, to provide further understanding of solid flammability in spacecraft environments.
Progress report on the deployment of the second Spacecraft Atmosphere Monitor Technology Demonstration Instrument
(2024 International Conference on Environmnetal Systems, 2024-07-21) Madzunkov, S.; Malone, C. P.; Simcic, J.; Jung-Kubiak, C.; Bae, B.; Kraus, H.; Nikolic, D.; Hristov, V.; Homer, M.L.; Fu, D.; Zhong, F.; Tillmans, T.; Garkanian, V.; Urrutia, J.; Lim, H.; Poliquin, E.; Kidd, R.D.; Nath, M.; Maiwald, F.; Hancock, B.; Darrach, M.
The second Technology Demonstration Unit (TDU2) of the Spacecraft Atmosphere Monitor (S.A.M.) is a miniaturized gas chromatograph mass spectrometer (GC/MS) instrument for monitoring the chemical composition of cabin air for human spaceflight missions. TDU2 was delivered to the International Space Station (ISS) on November 9, 2023, with the SpaceX 29th commercial resupply services mission (CRX-29). TDU2 will continuously monitor the major atmospheric constituents (MCA) and trace organic volatiles (TGA) with relative abundances ranging from parts per billion (ppb) to parts per million (ppm). The TDU2 uses the quadrupole ion trap mass spectrometer (QITMS) sensor and includes a micro-electromechanical system (MEMS) preconcentrator, gas chromatograph, and microvalve system. The ruggedized miniature form factor allows the aisle deployment of TDU2 to monitor cabin air quality in different locations and during various crew activities. We will discuss the operational performance of TDU2 and report on its first onboard measurements of more than twenty TGA compounds identified as harmful to cabin crew during long exposures.
3D-Printed Polymer-Based Conformal Space Radiation Shield with Heat Dissipation
(2024 International Conference on Environmnetal Systems, 2024-07-21) Xiao, Yue; Zou, An; Carlson, David; Rahman, Aminur; Zhang, Pu
With the rapid development of SmallSats/CubeSats and space computing, the current aluminum (Al) bulk shielding technology poses challenges in Size, Weight, and Power (SWaP) requirements. The bulk shielding thickness is determined by the electronics with the least radiation resistance, which prohibits wider adoption of Commercial-off-the-shelf (COTS) electronics and results in increased system weight. In addition, the absence of advanced thermal management technologies for SmallSats/CubeSats largely limits the adoption of high-power, low-price, and readily available COTS electronics. To address these challenges, we developed an innovative 3D-printed, lightweight polymer composite radiation shield comprising of a Metal Oxide Polymer Composite (MOPC) layer and a Fiber Reinforced Polymer Composite (FRPC) layer that is compatible with other polymer-based radiation shielding composites. Specifically, the MOPC utilizes high-atomic-number metal oxides that are infused into the base polymer to achieve superior radiation attenuation performance than Al. FRPC utilizes aligned carbon fibers in the same base polymer to achieve high in-plane thermal conductivity for electronics heat dissipation. In addition, such a shield is manufactured by the advanced five-axis Direct Ink Writing (DIW) 3D printer for further improved thermal transport and structural strength. Further, the 3D-printed shield enables locally enhanced spot shielding to provide additional radiation shielding for sensitive electronics without increasing the overall shield thickness. For MOPC, we have achieved an up to 7.4x increase in photon mass attenuation coefficient compared with Al and 18%–26% weight reduction for electron attenuation compared with Al at GEO. For FRPC, we achieved 2.55 W/m-K in-plane thermal conductivity, which is ~4x higher than the base polymer. Overall, the shield prototype can achieve 7° C reduction in CPU temperature or a 22.5% increase in CPU power, and if spot shielding is utilized, an estimated 62% weight reduction for 1U CubeSat systems can be achieved.
Radiometric Level Measurement (RLM)
(2024 International Conference on Environmnetal Systems, 2024-07-21) Flynn, Michael; Jang, Jeoungjae; Rodriguez-Rolon, Alondra S.
Conventional tank level sensors do not work in microgravity. The reason for this is that in microgravity water in a tank exists as a random mixture of droplets, films, saturated vapor and bulk water. This paper provides a preliminary analysis of a method to measure water in a tank in microgravity using Galactic Cosmic Radiation (GCR). GCR is charged high energy particles that come from random directions in space. When they collide with other molecules, such as the contents of a tank, they create secondary particles. If GCR encounters water, protons are produced as secondary particles. The number of protons produced is a function of the amount of water and the flux and energy of the GCR particles. In this study we use a technique called radiometric level measurement (RLM) to determine the amount of water in the tank by measuring the number of protons generated by GCR. RLM uses two proton sensors and one GCR sensor. One proton sensor is placed adjacent to the tank to measure proton generation. Another proton sensor is placed sufficient distance away from the tank to measure the background level of protons in space. A GCR sensor is also placed adjacent to the tank to measures the variation of the GCR background. The difference between the two proton sensors is a measurement of the water content of the tank at any specific GCR background level
Performance of a Regenerable Carbon Filter for the Plasma Pyrolysis Assembly
(2024 International Conference on Environmnetal Systems, 2024-07-21) Berger, Gordon; Agui, Juan; Mehan, Jeff; Crawford, Kagen
Oxygen recovery will be a key element of advanced Environmental Control and Life Support Systems (ECLSS) on future deep space missions. The Plasma Pyrolysis Assembly (PPA) works in conjunction with a Sabatier Reactor, which is a leading oxygen recovery technology. The function of the PPA is to recover hydrogen from the Sabatier reactor products. The 3rd generation PPA processes the methane, produced by the Sabatier Reactor, at a 4 crew-member flow rate to produce hydrogen, acetylene, other trace gases and solid carbon fines. A Regenerable Carbon Filter (RCF) was developed under a SBIR by Umpqua Research Company to capture the nuisance carbon fines and was put through integrated ground testing with a 3rd generation PPA unit at the NASA Marshall Space Flight Center's ECLSS Environmental Chamber (E-Chamber). The filter system consists of three separate in-series components: a regenerable electrostatic precipitator, a regenerable fibrous media filter, and a passive HEPA filter. Oxidation is used to regenerate the two regenerable components by flowing a small amount of oxygen through the regenerable components at 750 C. The tests consisted of carbon loading from the PPA and regeneration stages. Two different duration carbon loading cycle were performed corresponded to a typical loading cycle of the PPA, 8 hours, before the PPA reactor requires regeneration, and a 20-hour loading cycle to test for performance under longer carbon build up. The effectiveness of regeneration was checked visually inside the electrostatic precipitator component through borescope inspection. In addition, the level of carbon oxidation during regeneration was monitored by measuring gaseous products with a gas chromatograph. The results of the two loading tests and multiple regeneration tests will be presented in this paper.
Swift BAT Loop Heat Pipe Thermal Performance After 19+ Years On Orbit
(2024 International Conference on Environmnetal Systems, 2024-07-21) Choi, Michael
Swift is a NASA Medium-Size Explorer mission. It was designed for two years. The spacecraft was launched successfully into an orbit of 600-km altitude and 20.69° inclination on November 20, 2004. Because of its high science return, the Swift mission operation has been extended. The Burst Alert Telescope (BAT) Detector Array Blocks are mounted to the bosses of eight ammonia constant conductance heat pipes (CCHPs) embedded in the detector array plate. Two propylene loop heat pipes (LHPs), numbered #0 and #1, transport heat from the CCHPs to a radiator, which is located on the shaded side of the observatory. The BAT LHP thermal control subsystem is robust, including redundancy of LHPs, CCHPs, heater circuits and heater controllers. The LHP compensation chamber (CC) heater controllers had failures in 2005, 2010 and 2015. There was a fix for each of the failures for thermal recovery. Presently the LHP thermal control subsystem is still operating after 19+ years on orbit. It still maintains the average DM XA1 temperature nearly steady at 20° C, despite the LHP #1 CC temperature has intermittent droops. The droops are likely caused by the heater controller channel for the variable conductance heat pipe (VCHP) that uses a LHP CC set point offset approach. The heater controller failures provide a lesson learned for future design of heater controllers. The BAT radiator temperature and LHP condenser temperatures have no significant changes over 19+ years on orbit. It implies the 8-9 mil thick AZW/LA-II white paint on the radiator has no significant degradation and the optical properties are adequate for long term. It was the right decision to decrease the paint thickness from 13 mil to 8-9 mil to assure good adhesion, despite it increased the beginning of life absorptance by 0.02.
Assessment of Performance Degradation of the Spacesuit Water Membrane Evaporator (SWME) Materials Due to Environmental Conditions
(2024 International Conference on Environmnetal Systems, 2024-07-21) Abney, Morgan; Wilson, Sara; Pitts, Ray; Petersen, Elspeth; Bhattacharyya, Dibakar; Lipscomb, Glenn; Reeder, James; Steele, John
The Spacesuit Water Membrane Evaporator (SWME) is a next-generation design to replace the sublimator in spacesuits for thermal control. The SWME design includes nearly 28,000 hollow fiber membranes epoxied into a titanium housing. Limited data is available to understand the risk of exposure of the SWME materials to nominal and off-nominal conditions including organic chemicals, inorganic chemicals, particulates, and microbes. This effort sought to empirically investigate the effects of changing membrane characteristics such as pore size, porosity, and surface energy on membrane performance and the effects of environmental conditions on those characteristics. Performance data is presented and the long-term impacts on SWME performance are discussed. Effects of broken fibers, epoxy breakthrough, crimped or crushed fibers, and changes in fiber surface energy are also explored.
Predicting the Microgravity Performance of Terrestrial Portable Fire Extinguishers
(2024 International Conference on Environmnetal Systems, 2024-07-21) Abney, Morgan; Weislogel, Mark; Harper, Susana; Juarez, Alfredo; Willard, Doug; Provin, Tim; Goetter, Chris; Simpkins, Patrick; Santamaria, Rudy; Chen, Yongkang; Dietrich, Daniel; Urban, David; Williams, Dave
Portable fire extinguishers (PFEs) are a key component of spacecraft emergency response systems. The International Space Station uses custom PFEs to meet the unique microgravity and enclosed space requirements of the vehicle. For future missions, terrestrial commercial off-the-shelf (COTS) PFEs may offer more economical solutions. Depending on the design and layout of the targeted locations and chosen commodity, however, reduced gravity may affect the performance of terrestrial PFEs (e.g., two-phase systems). To better characterize the potential of using COTS hardware, two terrestrial PFEs, one charged with HFC-227ea pressurized with nitrogen gas and one charged with carbon dioxide, were modeled to predict performance in both 1-g and microgravity environments. Testing was conducted in 1-g in best and worst case configuration to validate the model. Here we provide a detailed description of the model, report the methods of PFE testing, and discuss the predicted effects of microgravity on PFE performance.

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