Design and Evaluation of Elements of a Life Support System For

Home , C change, Flight suit, Jessica Meir, Manned Maneuvering Unit, Pressure suit

Design and Evaluation of Elements of a Life Support System for Mechanical Counterpressure Spacesuits by Jeremy Paul Stroming Submitted to the Department of Aeronautics and Astronautics in partial fulﬁllment of the requirements for the degree of Master of Science in Aeronautics and Astronautics at the MASSACHUSETTS INSTITUTE OF TECHNOLOGY May 2020

Author...... Department of Aeronautics and Astronautics May 19, 2020

Certiﬁed by...... Dava J. Newman Apollo Professor of Aeronautics and Astronautics Thesis Supervisor

Accepted by ...... Sertac Karaman Chair, Graduate Program Committee 2 Design and Evaluation of Elements of a Life Support System for Mechanical Counterpressure Spacesuits by Jeremy Paul Stroming

Submitted to the Department of Aeronautics and Astronautics on May 19, 2020, in partial fulﬁllment of the requirements for the degree of Master of Science in Aeronautics and Astronautics

Abstract Mechanical counterpressure (MCP) spacesuits offer several advantages over tradi- tional gas-pressurized suits including lower energy cost of transport, reduced risk of decompression due to suit tear or puncture, and increased astronaut comfort. The BioSuitTM is an MCP concept being developed at MIT primarily for planetary extravehicular activity (EVA) on the Moon and Mars. In this thesis we present the initial design and testing of several key parts of the life support system for the BioSuitTM. First, a breathing compensation bladder that covers the chest and is pneumatically continuous with the helmet is discussed. This bladder eases the burden of respiration associated with constrictive garments and adapts to the changing volume of the chest to provide equalized pressure across the torso. The initial results of laboratory testing of the airflow and pressure maintenance of a bladder-helmet system on a mannequin are presented. Second, thermal modeling of a BioSuitTM EVA on the lunar and Mar- tian surfaces was conducted to assess the performance of MCP spacesuit garments in protecting astronauts from the extreme temperatures and harsh radiation environments of those locations. This modeling included new proposed radiation protection and insulating materials as well as a passive elastic compressive layer. Results were computed for both male and female astronauts, helping to identify suit design differences that will be needed to accommodate both men and women who will conduct future EVAs. This work is used to inform future design requirements for the suit’s thermal management system. Overall, this research advances the development of life support systems for a full MCP spacesuit and lessons learned can be applied for future engineering prototypes.

Thesis Supervisor: Dava J. Newman Title: Apollo Professor of Aeronautics and Astronautics

3 4 Acknowledgments

While an astronaut wearing a specially designed spacesuit can survive for short periods in a vacuum, a graduate student at MIT most deﬁnitely cannot. There are dozens of people who have helped, supported, and guided me during my two years in this program and throughout my six total years at MIT. Without them I could never have gotten to this point. As I write this in the midst of the unprecedented COVID-19 global pandemic and a socially distanced world, the critical importance of human support networks is more apparent than ever.

To Dava Newman, thank you for giving me the opportunity to stay in my home community of MIT AeroAstro and join you for two incredible years as a graduate student. Your skill in connecting students with incredible and far ﬂung opportunities is unparalleled. Your ability to manage a ceaselessly frenetic work schedule while maintaining an unwavering and sincere personal interest in your students, friends, and colleagues is something to be admired. My life changed forever when you responded to one shot-in-the-dark email I sent you during my senior year of undergrad and I joined you for a UROP.

For their invaluable assistance on the breathing bladder project, thank you to Gui Trotti and Michal Kracik. It would not have come together without you. To En- rico Rossetto, Vittorio Cafaggi, Nicola Parise, Andrea Azzolin, Alessandro Guzzon and all others at the D-Air LabTM who helped turn some of my sketches into a testable piece of hardware, grazie mille. Your hospitality during the week I spent in Vicenza was outstanding. It was far too short a visit and I hope to return someday.

To my friends in the Human Systems Lab, thank you for creating and sustaining one of the best communities in the AeroAstro department. Thank you especially to Tim McGrath for being such a fun and smart research partner and mentor. I’ll always remember our experience on that Zero G ﬂight. Thank you also to Golda, Allison, Alvin, Ferrous, Rachel, Shea, Becca and Maya for building such a great cohort and Aditi and Tom for providing sublime oﬃce banter. To the other

5 members of the Quals study group including Jess, Nick, Christine and Akshay, I’m so glad I had you with me to go through the experience. Thank you also to the HSL and AeroAstro staff including Quentin Alexander, Liz Zotos, Leah Lovgren, Raina Puels, Beata Shuster and Joyce Light for keeping everything together and always being there to help (or just share some gossip). And thank you to Jennifer Craig for teaching me to be a better writer and helping guide our 16.83 students all the way to Enceladus. To the NASA Space Technology Research Fellowship office, thank you for believing in me and funding my proposal. You granted me the gift of freedom and flexibility in my graduate career that not many students have the privilege to experience. Thank you also to my NSTRF mentor, Chris Massina, for orchestrating a visit for me at the Johnson Space Center. John and Jodi Graf have a special place in my heart for letting me stay in their home during my visit. Thanks to Jon Michael Tucker for being a welcoming and kind office mate. Even in just 3 weeks you shared lessons on humility and responsibility in engineering (and life) that I still think about. And thank you especially to Hee Jong Song who connected me with a student license of TAIThermTM, without which a chapter of this thesis could not have been written. Obrigado to Ligia Fonseca Coelho, Manuel Almeida and the MIT Portugal Program for bringing us together in the opportunity to build a NanoLab payload that flew to space. It was wonderful to get to know you over Skype for so many months before you finally visited me in Boston. I can think of no better friends to co-star in the miniature feature film they made of our project. While I won’t get to visit you in Portugal this summer, it is definitely on my list once the world returns to normal. Thank you to all of my different housemates. To my friends Andres, David, and Julius, I’m so happy we kept our close ties from undergrad alive. And to Althea, thank you for always being my biggest cheerleader and teaching me so much about everything ranging from Navajo culture to baby care. I’ll never forget our crazy, spontaneous adventure in Iceland. To my friends on the MIT Cycling team, I cannot thank you enough. Being a part of this club was absolutely the best part of my graduate school experience

6 and what I will remember and cherish the most. Thank you to Sarah for being an incredible co-captain. And to Nic, Miles, Berk, Joanna, Dmitro, Carolyn, Lee, Meia, Amy, Tori, Guillaume, Liam and all the others, thank you for all the rides and camaraderie, from coﬀee stops at Haute and RSC to big races like GMSR and marathon rides like Six Gaps and Fall and Winter Training Camps. I look forward to rolling with you all again soon. Finally, thank you to my family. My parents, Susan and Steve, deserve national recognition for their constant love and support from all the way across the country in Issaquah throughout my six years at MIT. My brother Ahren and my sister Signe have always been there for me to lean on, whether it’s to share a laugh or oﬀer a patient ear anytime I am struggling. Their own incredible achievements–saving the world from Puget Sound to India—inspire me every day. I love you all.

7 8 Contents

1 Introduction 19 1.1 Motivation ...... 19 1.2 Problem Statement ...... 21 1.3 Contributions ...... 22 1.3.1 BioSuitTM Breathing Bladder ...... 22 1.3.2 MCP Suit EVA Thermal Modeling ...... 22 1.4 Thesis Outline ...... 23

2 Background on Spacesuit Life Support Systems 25 2.1 Planetary EVA Environments ...... 25 2.2 Life Support System Requirements ...... 26 2.3 Historical EVA Suits ...... 29 2.3.1 Mercury and Gemini ...... 29 2.3.2 Apollo ...... 31 2.3.3 Shuttle and ISS EMU ...... 32 2.3.4 xEMU ...... 33 2.3.5 Summary ...... 35 2.4 Life Support Subsystems ...... 35 2.4.1 Atmospheric Revitalization ...... 35 2.4.2 Carbon Dioxide Scrubbing ...... 37 2.4.3 Thermal Management ...... 39 2.4.4 New and Advanced Thermal Management Technologies . . . . 44 2.4.5 Summary ...... 49

9 2.5 Mechanical Counterpressure Spacesuits ...... 49 2.5.1 Space Activity Suit ...... 50 2.5.2 MIT BioSuitTM ...... 54 2.6 Summary ...... 56

3 A Breathing Compensation Bladder for the BioSuitTM 59 3.1 Background ...... 59 3.1.1 Partial Pressure Flight Suits ...... 60 3.1.2 The Space Activity Suit ...... 63 3.1.3 MIT BioSuitTM ...... 67 3.2 Design and Analysis ...... 68 3.2.1 Physiological Requirements ...... 68 3.2.2 Design ...... 70 3.2.3 Helmet and Breathing Bladder Air Flow Analysis ...... 72 3.3 Fabrication ...... 75 3.4 Testing and Initial Results ...... 78 3.4.1 Individual Pressure Sensors Test ...... 81 3.4.2 Mat Sensor Test ...... 82 3.4.3 Discussion ...... 84

4 Thermal Modeling of BioSuitTM EVA on the Moon and Mars 87 4.1 Evaporative Cooling in an MCP Suit ...... 87 4.2 EVA Thermal Modeling ...... 90 4.3 Using TAIThermTM ...... 92 4.3.1 Human Model ...... 92 4.3.2 Spacesuit Model ...... 93 4.3.3 Weather Model ...... 96 4.4 Results ...... 99 4.4.1 Moon Modeling ...... 101 4.4.2 Mars Modeling ...... 106 4.5 Discussion ...... 112

10 5 Conclusions 117 5.1 Contributions ...... 118 5.1.1 BioSuit Breathing BladderTM ...... 118 5.1.2 MCP EVA Thermal Modeling ...... 118 5.2 Challenges and Limitations ...... 119 5.3 Future Work ...... 120 5.4 Final Conclusions ...... 122

A TAIThermTM EVA Human Visualizations 123 A.1 Moon EVA Simulation Results ...... 124 A.2 Mars EVA Simulation Results ...... 127

B Eﬀect of Layering a Garment in TAIThermTM 131

C Digital Archive 133 C.1 Contents of Digital Archive ...... 134

11 12 List of Figures

2-1 Evolution of NASA EVA spacesuits ...... 29 2-2 Ed White conducting the ﬁrst American EVA ...... 30 2-3 Apollo Portable Life Support System ...... 31 2-4 Space Shuttle EMU PLSS ...... 33 2-5 Schematic of EMU Primary Oxygen Ventilation Loop ...... 36

2-6 Schematic of RCA CO2 Removal System ...... 39 2-7 Human Metabolic Heat Output at Diﬀering Activity Levels ...... 40 2-8 Liquid Cooling Ventilation Garment ...... 41 2-9 Schematic of Ventilation in the EMU ...... 42 2-10 EMU Sublimator ...... 43 2-11 Schematic of the LCAR Absorbtion Process ...... 46 2-12 Comparison of Heat Rejection using the SWME and SEAR ...... 47 2-13 The Space Activity Suit and Breathing Bladder ...... 51 2-14 Subject Mobility Testing with the Space Activity Suit ...... 52 2-15 Space Activity Suit Life Support System Diagram ...... 53 2-16 Mock-up of the MIT BioSuitTM ...... 54 2-17 Renderings of the BioSuitTM and PLSS Backpack ...... 55 2-18 Rendering of a Proposed BioSuitTM Breathing Bladder Design . . . . 56

3-1 S-1 Partial Pressure Flight Suit ...... 61 3-2 MC-1 Partial Pressure Flight Suit ...... 62 3-3 MC-3 Partial Pressure Flight Suit ...... 63 3-4 The Space Activity Suit ...... 64

13 3-5 Helmet and Bladder Diagram for the Space Activity Suit ...... 65 3-6 Diagram of Space Activity Suit Breathing Bladder ...... 66 3-7 BioSuitTM Rendering and Breathing Bladder ...... 67 3-8 Lung Tidal Volumes for Different Breathing Types ...... 69 3-9 CAD Model of Breathing Bladder ...... 70 3-10 CAD Modeling of a Considered Air Hose Bladder Design ...... 71 3-11 Using ExactFlat to Produce Drawings for Bladder Manufacturing . . 72 3-12 Flow Model of Breathing Bladder ...... 73 3-13 Pressure Contour Map from Bladder Flow Modeling ...... 74 3-14 BioSuitTM Bladder Modeling at the Dainese D-Air LabTM ...... 75 3-15 Breathing Bladder Prototype at DaineseTM ...... 76 3-16 BioSuitTM Breathing Bladder Prototype ...... 77 3-17 Schematic of Test Setup for Bladder Inflation ...... 79 3-18 Pressure Sensor Setup for Bladder Testing ...... 80 3-19 Time Lapse of Bladder Inflation ...... 81 3-20 Plot of Individual Sensor Pressure Response ...... 82 3-21 Final Pressure Sensor Configuration ...... 83 3-22 Before and After View of novel pliance® User Interface During Testing 84 3-23 Plot of Mat Sensor Pressure Response ...... 84

4-1 Layering Comparison of the EMU and the BioSuitTM ...... 88 4-2 Body Parts of the TAIThermTM Human Model ...... 91 4-3 TAIThermTM Human Model Overview ...... 92 4-4 EVA Metabolic Proﬁle ...... 93 4-5 TAIThermTM Garment Layup ...... 94 4-6 Boron Nitrite Nanotube and Aerogel Material Test Coupons . . . . . 95 4-7 Lunar and Martian Surfaces ...... 97 4-8 Martian Weather Model ...... 98 4-9 Thermal Results for Lunar EVA Wearing Skinsuit ...... 102 4-10 Skin and Core Temperatures for Lunar EVA ...... 103

14 4-11 Heat Fluxes and Sweat Rate for Lunar EVA ...... 104 4-12 Skin and Core Temperature Comparison for Male and Female Astro- nauts During Lunar EVA ...... 106 4-13 Heat Fluxes and Sweat Rate for a Female Astronaut During Lunar EVA107 4-14 Thermal Results for Mars EVA Wearing Skinsuit ...... 108 4-15 Skin and Core Temperatures for Martian EVA ...... 109 4-16 Thermal Results for Mars EVA Wearing Aerogel Insulation Layer . . 110 4-17 Heat Fluxes and Sweat Rate for a Male Astronaut During Martian EVA111 4-18 Skin and Core Temperature Comparison for Male and Female Astro- nauts During Martian EVA ...... 112 4-19 Heat Fluxes and Sweat Rate for a Female Astronaut During Martian EVA...... 113

A-1 Thermal Results for Lunar EVA Wearing Skinsuit ...... 124 A-2 Thermal Results for Lunar EVA Wearing Skinsuit with BNNT jacket 125 A-3 Thermal Results for Lunar EVA Wearing Aerogel Insulation Layer . . 126 A-4 Thermal Results for Mars EVA Wearing Skinsuit ...... 127 A-5 Thermal Results for Mars EVA Wearing BNNT with Skinsuit . . . . 128 A-6 Thermal Results for Mars EVA Wearing Aerogel Insulation Layer . . 129

B-1 Layering Comparison Test ...... 132

15 16 List of Tables

2.1 General Spacesuit Life Support System Requirements ...... 27 2.2 EMU PLSS Component Mass Breakdown ...... 34

3.1 Overview of Historical Applications of Volume Compensatory Breath- ing Bladders in Aerospace ...... 60 3.2 Statistical Summary of novel pliance® Pressure Data ...... 85

4.1 Comparison of metabolic cooling systems for the EMU, SAS and BioSuit 89 4.2 MCP Suit Material Properties ...... 94 4.3 Moon and Mars Weather Comparison ...... 97 4.4 TAIThermTM Modeling Testing Matrix ...... 99 4.5 Core and Skin Temperatures for a Range of Human Thermal States . 100 4.6 Summary of TAIThermTM EVA Modeling Results ...... 114

17 18 Chapter 1

Introduction

1.1 Motivation

When cosmonaut Alexei Leonov became the first person to venture outside of a space capsule on March 18, 1965, it marked to first demonstration of an Extravehicular Activity (EVA) suit. Though Leonov remained tethered to his Voskhod 2 spacecraft with an air supply umbilical, this was the first test of a life support system used to support an astronaut outside of a spacecraft. It did not go smoothly. Leonov’s pressurized suit expanded so much in the ambient vacuum that it became almost immovably stiff. After 12 minutes out, he got stuck trying to reenter the airlock, vented air to depressurize the suit and became overheated from exertion. The Soviet Union did not attempt another EVA for almost four years [1]. US astronaut Ed White followed Leonov three months later to complete NASA’s first EVA. White was tethered to the Gemini capsule through a life support umbilical as well. Though he experienced fewer mishaps than Leonov, White also had mobility difficulties and quickly became tired and overheated. The design problems in these early spacesuits extended beyond just the life support systems, but illustrate the importance of a robust air supply and thermal management system for the health, safety, and productivity of astronauts. Spacesuits engineered for subsequent programs including Apollo, Shuttle and the International Space Station (ISS) have improved upon the life support equipment

19 used in early spacewalks and have built up an impressive flight heritage. Still, the Portable Life Support System (PLSS) of the Extravehicular Mobility Unity (EMU)— NASA’s current spacesuit used on the ISS—is a product of incremental adaptations and modifications to the system engineered to support the astronauts of the Apollo program. The retirement of the Shuttle and the announcement of an intended return to the Moon in the late 2020s through the Artemis program offers an opportunity to examine new advanced technologies that could become a pivotal aspect of future human space exploration, improving efficiency, reliability and safety [2]. NASA has recognized the need for advanced spacesuit life support equipment as the agency plans not only a return to the Moon but also missions to Mars in the 2030s. NASA’s 2020 Technology Taxonomy and 2015 Space Technology Roadmap both highlighted the need for more advanced suit and life support systems:

“Advances in the pressure garment for future spacesuits are sought to allow efficient donning or doffing, enable long-duration and high performance suited operations in both space and planetary environs, provide scaling across a range of anthropometric sizes, and provide crew support, such as safe in-suit waste management and provision of water...Other PGS advancements include advanced and multi-functional suit materials and fabrication techniques; suit configurations and architectures with decreased mass; components with increased capability (e.g., mobility); sizing to accommodate a wide range of crew members; dust resistance technologies; and increased durability or life.”

“Advancements in the PLSS are necessary to enable future missions, such as to planetary environments. An example is the sublimator currently utilized as a part of the ISS Extravehicular Mobility Unit (EMU) to reject heat (metabolic heat plus system-generated heat); it sublimates water and the vapor is released into space. Sublimation is a physical mechanism requiring a hard vacuum environment, and thus cannot be used for Mars surface missions. Another example is the current ISS EMU, which uses

20 either a lithium hydroxide (LiOH) or silver oxide (MetOx) single sorbent

bed to absorb CO2 during an EVA. The LiOH is relatively light but not reusable and therefore has a signiﬁcant logistical and system mass impact for longer missions. The MetOx is regenerable between EVAs but has a mass greater than 18 kilograms (kg) when scaled for surface exploration metabolic loading, and also requires signiﬁcant power and vehicle space for the hardware used for regeneration. For Mars missions, the PLSS design will be pursued as closed-loop to function in the Mars atmosphere and be compliant with planetary protection guidelines and policy” [3, 4].

In addition to government agencies, private companies also aim to establish human space operations in the years ahead, meaning more efficient and reliable spacesuit and life support technologies will be needed. Mechanical counterpressure (MCP) spacesuit designs including the MIT BioSuitTM offer many potential advantages for future EVAs on the Moon and Mars including decreased mass, improved mobility, abrasion resistance and increased thermal operational range. However, this suit architecture is a fundamental departure from previous suit systems, requiring a redesigned life support system. Much research has been performed to investigate the physiology of pressure suits, the materials needed for MCP, and donning and doffing concepts, but there has been very little investigation into the elements needed for a BioSuitTM life support system. This thesis aims to fill several of these gaps.

1.2 Problem Statement

There remains a lack of knowledge as to the requirements of a life support system for an MCP spacesuit as well as what elements are needed and what methods are suﬃcient to provide pressure, oxygen, water and thermal regulation for an astronaut during EVA. The primary goal of this thesis is to develop prototypes of MCP life support system elements—speciﬁcally a torso breathing bladder—and also to investigate methods of thermoregulation within an MCP suit by conducting computational thermal modeling of planetary EVAs.

21 1.3 Contributions

The contributions of this thesis work can be broken down into two categories: the design, fabrication and testing of a prototype compensatory torso breathing bladder

for the BioSuitTM and a computational thermal analysis of EVA in lunar and Martian conditions to simulate perspiration-based cooling for an MCP suit. This study also included an assessment of the performance of several advanced suit materials and the thermoregulatory diﬀerences between female and male astronauts.

1.3.1 BioSuitTM Breathing Bladder

A prototype of a volume compensatory torso breathing bladder was design, fabricated, and tested in a laboratory setting. A bladder connected to the helmet airspace is necessary to accommodate the changing torso volume during respiration and equally distribute contact pressure across the chest. This work showed that modern analysis and manufacturing processes can expedite the development of the BioSuitTM and its subsystems. The bladder prototype advances the development of the life support system for the BioSuitTM as a whole and can now serve as a baseline for future improvements.

1.3.2 MCP Suit EVA Thermal Modeling

Detailed thermal modeling of a four hour astronaut EVA on the lunar and Martian surfaces was conducted. The model included a human wearing clothing layers representative of a mechanical counterpressure spacesuit such as the BioSuitTM. The simulations assessed the feasibility of relying on perspiration and evaporative cooling to provide thermal control for an MCP spacesuit. The overall thermal performance of diﬀerent advanced materials with potential applications in suit radiation shielding and insulation were also investigated as well as diﬀerences in thermal response between male and female astronauts. These simulations used an industry standard software called TAIThermTM with a built in human model more detailed than previous thermal analyses conducted.

22 1.4 Thesis Outline

The original contributions of this thesis are presented in two chapters. The two contributions represent different elements of a life support system for the BioSuitTM. Chapter 3 focuses on pressure regulation and breathing enabled by a chest bladder. Chapter 4 focuses on thermal requirements and temperature regulation of the suit during planetary EVA by leveraging cooling methods including perspiration that are not available in current gas pressurized suits. The objective of Chapter 2 is to review the history of the Portable Life Support System in American EVA spacesuits from the Gemini program all the way up to the xEMU suit currently in development at NASA. The chapter also reviews the subsystems that are part of the complete life support system for a spacesuit and summarizes several current proposals for new life support technology to be introduced into future EVA suits. Finally, mechanical counterpressure suits and research in their life support systems are reviewed. Chapter 3 discusses the history of compensatory torso breathing bladders used in aerospace applications and then details the design, fabrication, and initial testing of a prototype breathing bladder for use in mechanical counterpressure suits. Chapter 4 presents the setup and results of a computer thermal modeling campaign conducted to assess the feasibility of relying on evaporative perspiration to provide thermoregulation in an MCP during EVAs on the Moon and Mars. The model inputs include human physiology, environmental conditions and suit material properties. Results including body core and skin temperatures, comfort and thermal sensation, heat fluxes and sweat rates are presented. The performance of advanced spacesuit materials as well as differences between male and female astronauts are discussed. Chapter 5 summarizes the contributions made by this thesis and provides conclusions from this work as well as recommendations for future work.

23 24 Chapter 2

Background on Spacesuit Life Support Systems

This chapter introduces the environmental challenges of planetary EVA and the requisite components of a Portable Life Support System (PLSS) needed to keep an astronaut alive. It will review the history of the PLSS through multiple suit generations and also highlight individual subsystems and components of the PLSS of NASA’s current spacesuit, the Extravehicular Mobility Unit (EMU). Proposals for future improvement are also discussed. Finally, this chapter discusses the development history and theoretical advantages of a mechanical counterpressure (MCP) spacesuit design. The life support systems of two MCP suits—the Space Activity Suit and the MIT BioSuitTM—are introduced.

2.1 Planetary EVA Environments

All spacesuit life support systems are designed to protect a human during operation in extreme environmental conditions. EVAs so far have only occurred in Low Earth Orbit (LEO) and on the surface of the Moon during the Apollo program. In the future it is likely that humans will again conduct EVAs on the Moon as well as journey to Mars for the ﬁrst time. A summary of the environmental conditions found in each of these three locations is provided here. Although local conditions on the Moon and

25 Mars can vary drastically, future spacesuits will likely be designed for a broad range of conditions that may be experienced on the surface. The LEO environment is characterized by microgravity and the hard vacuum of space. Space itself has no temperature, but objects in orbit can vary from -100°C in the dark to +136°C in sunlight. The exterior surfaces of the EMU can get as hot as 90°C in direct sunlight, requiring a 20-minute airlock cool-down before an astronaut may remove the suit and return to the pressurized spacecraft [5]. At the altitude of the ISS (409 km), the day-night cycle lasts approximately 90 minutes.

1 The lunar surface is characterized by approximately 6 (0.17) Earth g, and surface temperatures ranging from -143°C (-225°F) to +127°C (+261°F) due to direct sunlight or shading in an ambient vacuum. In direct sunlight, visible and ultraviolet light can be re-radiated from the lunar regolith in the infrared (IR) spectrum, amplifying the heating effect. In deep shadowed craters or during the extended night of the lunar 28-day day-night cycle, temperatures can approach absolute zero Kelvin (-273°C) [6]. Temperatures are also dependent on latitude and local lighting conditions. During a lunar EVA, a crew member will experience sharp transient effects as they move in and out of shadows and change their orientation with respect to the sun [7]. The Martian surface is characterized by 0.38 Earth g, atmospheric temperatures from -143°C (-225°F) to +20°C (+68°F) with seasonal variations, and a thin atmosphere of 0.6-1.0 kPa composed of 95% CO2 and wind gusts of up to 30 m/s that add convective heat loss. Ground temperatures range from -118°C (-180°F) to +29°C (+85°F). Mars has a 24.6 hour day-night cycle, and sunlight is about 43% the in- tensity experienced on Earth [6]. Overall, conditions vary dramatically depending on specific location, time of the year, and dust content in the atmosphere [8].

2.2 Life Support System Requirements

As a baseline, a human has several key requirements for survival on the order of 8 hours needed for an EVA:

• Oxygen for breathing

26 • Removal of carbon dioxide

• Pressurized atmosphere

• Temperature control

These requirements must be provided in order to protect the human from the extreme environments described in the previous section. In the EMU, these four functionalities are provided by the PLSS. Beyond these essentials, the PLSS also provides power for the suits electronic components, enables radio communication capabilities and monitors biometric data [5, 9]. Table 2.1 summarizes the basic requirements that every spacesuit life support system must provide to keep the wearer alive for the duration of an EVA in a space environment.

Table 2.1: General Spacesuit Life Support System Requirements

Life Support Subsystem Functional Requirements Minimum 20.7 kPa (3 psi) oxygen supply for 8 hours; Oxygen Ventilation minimum of ∼0.25 kg (0.55 lbs) total oxygen for 8 hours Ventilation ﬂow rate of ∼6.0 ft3/min (0.17 m3/min); Carbon Dioxide Removal partial pressure of CO2 nominally less than 0.5 kPa (0.07 psi) Pressurization Minimum of 20.7 kPa (3 psi) absolute applied pressure to the body Suit internal temperatures 18-27 C (65-81 F); Temperature Control ° ° humidity 30-70%

At sea level on Earth, the atmosphere evenly exerts a total pressure of 14.7 psi (101 kPa). This air is composed of approximately 21% oxygen, 78% nitrogen, and 1% trace elements. Each gas exerts a partial pressure proportionate to its concentration in the atmosphere. Therefore, oxygen has a partial pressure at sea level of 3.1 psi (21.4 kPa). Humans require a minimum of about 3 psi (20.7 kPa) ambient oxygen for comfortable respiration although some people on Earth live at higher elevations where the partial pressure of oxygen is slightly lower. The true minimum is about 2 psi (14.5 kPa), the point at which the partial pressure of oxygen in the alveoli in the lungs matches the oxygen pressure in the capillaries and gas exchange will cease to occur. Past EVA suits have all used a pure oxygen environment with pressures ranging from 3-6 psi (20.6-41.4 kPa) as a compromise to balance interfering with suit mobility and minimizing the decompression time required when the suit is donned.

27 To provide an adequate margin of safety, a minimum partial pressure of oxygen of 3.5 psi (24.1 kPa) is recommended [10]. For example, the Apollo suit operated at 3.7 psi, but the modern Shuttle and ISS EMU operates at 4.3 psi (29.6 kPa) while the Russian Orlan suit operates at 5.8 psi (40 kPa) [5].

This oxygen is distributed throughout the suit by the ventilation system. This system is important for two reasons. First, gas flow helps cool the human body through convection, and second, some ventilation is required in any closed volume to keep waste gases such as carbon dioxide and odors at a healthy and tolerable level. Suit flow rates are generally between 4.5 and 6.0 cubic feet (0.13 to 0.17 cubic meters) per minute (cfm). The Apollo suit operated at 5.5 cfm (0.16 m3/min) and the EMU operates at a minimum of 6.0 cfm (0.17 m3/min) gas flow into the helmet. This flow is enough to prevent the buildup of carbon dioxide (hypercarbia) which can disrupt oxygen exchange in the lungs and cause headaches, dizziness, cognitive impairment and unconsciousness. Ventilation also keeps the temperature and humidity inside the suit within the narrow range acceptable for human comfort. Temperatures range from 18-27 degrees Celsius (65-81 Fahrenheit) with the setpoint being 22°C (72°F). Humidity ranges between 30-70% depending on the activity level [5, 11].

In the EMU, carbon dioxide is removed by the Contaminant Control Cartridge (CCC). This is either a lithium hydroxide filter bed that can be recharged on the ground or a metal oxide (Metox) filter that can be regenerated on the ISS. The CCC is sized to remove 0.67 kg (1.48 lbs) of CO2 before replacement. This is equivalent to a metabolic rate of 249 Watts per hour for 8 hours. At metabolic rates below 468 W, carbon dioxide is kept at a partial pressure less than 0.147 psi (1.0 kPa) and below 0.29 psi for metabolic rates between 468 and 586 W. During nominal operations it is desirable to keep the partial pressure of CO2 less than 0.07 psi (0.5 kPa) to prevent any of the effects of hypercarbia [5, 11].

28 2.3 Historical EVA Suits

EVA suits are unique in the realm of extreme environment protective gear in the need for a fully integrated life support system. While the history of American EVA suits would seem to linearly progress from program to program, this is not necessarily the case. Different contractors have been hired by NASA for different programs, including B.F. Goodrich for the Mercury suit, the David Clark Company for Gemini, and ILC Dover and Hamilton Standard for Apollo. Additionally, dozens more suit prototypes have been developed but never operated in the vacuum of space [5]. However, it is still useful to trace the evolvable aspects of the life support system by examining the space suits built for different EVA campaigns over the past 60 years. Figure 2-1 shows the evolution of EVA suits used by NASA from the Gemini program through the planned Artemis program.

Figure 2-1: An evolution of NASA EVA spacesuits. From left to right is the Gemini G4C suit worn by Ed White in 1965, the Apollo A7LB suit worn by Eugene Cernan in 1972, the Space Shutte EMU worn by Franklin Chang-Diaz in 2002, the ISS EMU worn by Jessica Meir in 2020, and a prototype xEMU suit modeled by Kristine Davis in 2019. All image credits to NASA.

2.3.1 Mercury and Gemini

When the Mercury program began in the late 1950s, NASA was in its infancy and had no suit programs of its own. It therefore had to rely on suits developed for high altitude aviation for the U.S. Navy and Air Force. The Mercury suit was a modiﬁed U.S. Navy Mark-IV high altitude ﬂight suit. The ventilation system was entirely

29 vehicle supported and included an air hose that entered at the hip and circulated ﬂow around the astronaut before exiting out of the helmet face. When the astronaut lowered the visor, the suit sealed, and air was vented through a small hose at the right side of the helmet [9].

Figure 2-2: Ed White wears the adapted Gemini G4C spacesuit during the ﬁrst American EVA on Gemini 4, June 3, 1965. The Ventilation Control Module can be seen on his chest attached to the gold life support umbilical, allowing the Gemini capsule to provide ventilation. Credit: NASA

Gemini spacesuits were modified U.S. Air Force X-15 high altitude flight suits. The Gemini G4C suit used in the first ever American EVA by Ed White included a Ventilation Control Module (VCM) mounted on the astronaut’s chest, as shown in Figure 2-2. During EVA outside the Gemini capsule, umbilical tethers provided oxygen and ventilation to the astronaut as they ventured outside the vehicle. All Gemini EVAs had vehicle supported life support systems. An Extravehicular Support Pack- age (ESP) and an Astronaut Mobility Unit (AMU) were developed to each provide one hour of untethered oxygen ventilation to the suit but were scrapped before they could be tested in flight. However, these modules formed the basis for later Apollo PLSS systems and the Manned Maneuvering Unit (MMU) propulsion system used in

30 the Shuttle EMU [9].

2.3.2 Apollo

The Apollo A7 suit was the ﬁrst to include a completely self-contained life support system, housed in the Primary Life Support System (PLSS) backpack. Engineers at Hamilton Standard built several components that would go on to become staples for the Shuttle and ISS EMU including the Liquid Cooling and Ventilation Garment (LCVG) and the feedwater sublimator for suit cooling.

Figure 2-3: The Apollo Portable Life Support System (PLSS) on display next to an LiOH (Lithium Hydroxide) CO2 scrubbing canister.

The Apollo PLSS weighed between 38 and 58 kg (85-129 lbs) depending on various conﬁgurations of which about 18 kg (40 lbs) was dedicated to the Oxygen Purge System (OPS), a secondary oxygen supply in case the primary ventilation system failed or the garment had a breach. It had enough ventilation capability to support a 4-hour EVA at an average metabolic rate of 351 Watts or a 6 hour EVA at an average

31 of 272 Watts [12]. Figure 2-3 shows the Apollo PLSS preserved at the Johnson Space Center without its protective cover. The OPS is removed but was mounted on top of the gold radio module. The Apollo suit was pressurized to 3.9 psi. For later missions including Apollo 15. 16 and 17, the PLSS capacity was increased to support 270 Watts of cooling for up to 8-hour EVAs and the nominal PLSS/OPS system mass correspondingly increased to 58 kgs (129 lbs) [12].

2.3.3 Shuttle and ISS EMU

The Shuttle EMU was redesigned from the A7 suit for free ﬂoating operation in LEO. The PLSS was increased in size to support EVAs lasting up to 8 hours at a metabolic rate of 249 W or 293 W for 7 hours. The primary components of the Shuttle PLSS are shown in Figure 2-4. The EMU can dissipate heat at a peak of 586 W for 15 minutes. It is pressurized to 4.3 psi (29.6 kPa) during EVA operations, requiring a 2.5 hour prebreathe to step down from the 14.7 psi (101.3 kPa) atmospheric cabin pressure in the ISS [11]. The PLSS is comprised of four circuits:

• Oxygen Ventilation Circuit

• Primary Oxygen Circuit

• Feedwater Circuit

• Liquid Transport Circuit

A Secondary Oxygen Pack (SOP) is attached to the bottom of the PLSS to provide 30 minutes of backup pressure and oxygen in case of an emergency but as of this writing it has never had to be used. Figure 2-4 shows an exploded schematic of the EMU PLSS. The total life support system has a mass of approximately 89 kg (196.6 lbs), of which 46 kg (101.4 lbs) is the PLSS, 6.5 kg (14.3 lbs) is the Display and Controls Module (DCM), 10.5 kg (23.1 lbs) is the SOP, 14.5 kg (32.0 lbs) is the

32 Figure 2-4: The Space Shuttle EMU PLSS [11]

Metox CCC, 6.7 kg (14.7 lbs) is the battery, 4.3 kg (9.5 lbs) is the radio, and 0.7 kg (1.6 lbs) is the antenna and cabling. The total rises to (205.2) lbs with the inclusion of the 3.9 kg (8.6 lb) wet LCVG [11]. This mass breakdown is shown in Table 2.2. A more detailed analysis of the EMU life support system is given in the following section.

2.3.4 xEMU

The Exploration EMU (xEMU) is NASA’s intended next generation spacesuit that will be used in the construction of the orbiting Gateway station around the Moon as well as the ﬁrst lunar EVAs as part of the Artemis Program. As of this writing, demonstration is scheduled for 2023 on the ISS before the lunar campaign envisioned for the late 2020s. Key changes include improved hip and shoulder mobility through rotating joint bearings and eventually the implementation of the Spacesuit Water Membrane Evaporator (SWME) to replace the legacy sublimator. It will also include

33 Table 2.2: EMU PLSS Component Mass Breakdown Component Name Mass Portable Life Support System (wet. with O2) 46.0 kg (101.4 lbs) Display and Controls Module (wet) 6.5 kg (14.3 lbs) Secondary Oxygen Pack (charged, with O2) 10.5 kg (23.1 lbs) Contaminant Control Cartridge (Metox) 14.5 kg (32.0 lbs) 2000 Series Battery 6.7 kg (14.7 lbs) Space to Space EMU Radio 4.3 kg (9.5 lbs) Antenna and Cable 0.7 kg (1.6 lbs) Liquid Cooling Ventilation Garment (wet) 3.9 kg (8.6 lb) Total 93.1 kg (205.2 lbs)

a dual swing bed Rapid Cycle Amine (RCA) CO2 scrubber. The RCA weighs less than the Metox canister and cleanses itself, meaning less servicing time, but is only operable in a vacuum. More detailed descriptions of both of these devices are given in later sections of this chapter. Oxygen ﬂow rate is targeted at the same set-point

3 of 6 cfm (0.17 m /min) to provide CO2 washout, but the fan is operable across the range of 4-7 cfm (0.11-0.2 m3/min) to accommodate testing of lower ﬂow rates [13].

The suit has a requirement to perform fully autonomously for up to 8 hours at a metabolic rate of 352 W with a peak load of 810 W. Additionally, the highly pressurized Secondary Oxygen Pack has been replaced by system that includes a Secondary Oxygen Loop, essentially installing a second oxygen cylinder to actively cycle on and oﬀ to assist the Primary Oxygen Loop when it is needed. Both loops are charged to the same pressure (3000 psi/20,600 kPa) which is much less than the dangerous and single use 6000 psi (41,300 kPa) SOP system that it replaces. Internal suit pressure can be increased up to 8.4 psi (57.9 kPa) to allow for a zero prebreathe time for certain vehicle conﬁgurations or provide emergency Decompression Sickness treatment [13].

34 2.3.5 Summary

This section reviewed the life support systems of historical and future NASA spacesuits including the Gemini G4C, Apollo A7L, Shuttle and ISS EMU and xEMU. The next section will take a deeper dive into the individual subsystems that together form the life support system of a spacesuit. The technology implemented in the EMU is used as a case study.

2.4 Life Support Subsystems

Now that an overview of past and present EVA suits has been conducted, we may take a closer look at several of the speciﬁc life support system functionalities including atmospheric revitalization, carbon dioxide scrubbing and thermal management. The technology used in the ISS EMU is described here as it represents the state of the art. The suit has been in continuous operation since the 1980s but receives periodic hardware refurbishments and upgrades.

2.4.1 Atmospheric Revitalization

Oxygen for breathing, air ventilation for cooling, and scrubbing for contaminant control are conducted by three loops within the EMU LSS: the Oxygen Ventilation Circuit, the Primary Oxygen Circuit and the Secondary Oxygen Pack. The Primary Oxygen Ventilation Loop delivers oxygen for breathing, provides pressurization for IVA and EVA activities, cools the interior of the suit, and removes exhaled gases and odors. Oxygen enters the suit through a manifold connecting the PLSS backpack to the Hard Upper Torso (HUT) where ducts direct the gas to the rear of the helmet. The oxygen ﬂows through the helmet vent pad and down in front

of the astronauts face, defogging the visor and pushing CO2 away from the mouth and nose. Air then circulates through the interior volume of the suit before being pumped back into several ventilation tubes located at the wrists and ankles. These tubes direct the gas back to the PLSS where it passes through the carbon dioxide

35 Figure 2-5: Schematic of the Shuttle EMU Primary Oxygen Ventilation Loop and components highlighted in yellow. In general, air flows into the suit from a vent pad and the back of the helmet and is returned to the PLSS via ventilation tubes in the LCVG. The oxygen has CO2 removed by a scrubber, cooling and humidification from the sublimator, fresh oxygen added by the tank and is pumped back into the suit again. Image from Reference [11]. scrubbing Contaminant Control Cartridge (CCC). Downstream of the CCC is the centrifugal fan that forces air through the system at the 6 cubic feet per minute (0.17 m3/min) standard flow rate by rotating at 19,300 rpm and maintaining a 0.126 psi pressure rise (0.869 kPa). Next, the gas travels through a heat exchanger that cools the air to about 13 degrees Celsius (55 Fahrenheit) and removes the humidity. A vent flow sensor precedes a regulator that adds oxygen from the primary or secondary oxygen tanks to the flow as needed. Just before the dry gas reenters the manifold and suit, a pressure transducer and a carbon dioxide sensor warn if the pressure is outside acceptable limits or if the scrubbers are not cleansing. If the scrubbers fail, expired gas can be vented through a one-way check valve into space and the oxygen tanks opened for an open loop breathing gas supply [11]. An original schematic of this system is shown in Figure 2-5.

36 The Primary Oxygen Circuit provides oxygen from storage tanks for breathing, as well as suit and water tank pressurization. Two bottles store a minimum of 0.55 kg of oxygen at about 900 psi (6200 kPa). This oxygen is siphoned oﬀ through a two stage regulator to step down the pressure into the ventilation loop. A separate regulator empties into the feedwater tank to pressurize the feedwater circuit used for thermal control [11].

The Secondary Oxygen Pack (SOP) provides ventilation backup in case of failure of the Primary Oxygen Circuit and also provides the capability of 30 minutes of open loop oxygen supply and purge in case of emergency. The SOP consists of two spherical bottles containing a total of 1.19 kg of oxygen pressurized to 6000 psi (42000 kPa). A two-stage regulator is used to siphon gas from the bottles into the suit if necessary [11].

2.4.2 Carbon Dioxide Scrubbing

Chemical processes have been used in all spacesuit life support systems for removing carbon dioxide from the breathing gas. Summarized below are three technologies used for CO2 scrubbing. Lithium Hydroxide (LiOH) cannisters are the oldest system and were deployed in the Apollo A7L suit as well as the early Shuttle EMU. They have since been replaced by Metox scrubbing cartridges. The next generation suit, the xEMU will use a new system called Rapid Cycle Amine that promises mass reductions and improvements in eﬃciency as compared to earlier technologies.

Lithium Hydroxide

Lithium Hydroxide (LiOH) has the longest ﬂight heritage for use as a carbon dioxide removal system, dating back to the Apollo program. Air is directed through canisters with perforations that allow air to react with the solid LiOH material inside. The carbon dioxide bonds to the lithium hydroxide forming solid lithium carbonate and water vapor. The system is governed by the following reaction:

37 2LiOH(solid) + CO2(gas) → Li2CO3(solid) + H2O(gas) (2.1)

In general, this reaction is non-hazardous, but canisters are single use and have a limited life. They have been phased out of the EMU since the Shuttle program.

A single 2 kg LiOH canister removes only one person’s average daily output of CO2 making it mass prohibitive for long duration missions[14].

Metox

Metal Oxide (Metox) cartridges have generally replaced LiOH for most EVA applications as these cartridges can be cleansed and reused. Once a cartridge in an EMU

PLSS is saturated, it can be taken aboard the ISS and heated to release the CO2. In this system, carbon dioxide adsorbs to a silver oxide solid in the cartridge, forming silver carbonate. The system is governed by the following reaction:

Ag2O(solid) + CO2(gas) → Ag2CO3(solid) (2.2)

The system oﬀers the beneﬁt of reusability, but recharge of EMU cartridges takes 8-10 hours at a temperature of 140°C in a 1 kW oven and then a 4-hour cooldown.

It also releases the CO2 back into the station where larger scrubbers must capture it again. Additionally, silver is very dense and the cartridges have a mass of 14.5 kg, ﬁve times as much as the LiOH canisters [14, 15].

Rapid Cycle Amine

The Rapid Cycle Amine (RCA) dual swing-bed technology is to replace the Metox cartridges for the new xEMU. Using a proprietary amine chemical developed by Hamilton Sundstrand (now Collins Aerospace). One sorbent bed is in series with the ventilation loop of the suit, adsorbing water vapor and carbon dioxide. The reaction is exothermic, generating heat that is transferred to the second bed. This second desorbing bed is exposed to the vacuum of space and where the carbon dioxide and water is vented to space in an endothermic reaction. Once the adsorbing bed is saturated,

38 a valve actuates to switch the desorbing bed back into the ventilation loop and the saturated bed is exposed to vacuum. Figure 2-6 shows a schematic of the swing-bed process. Note the piston that changes which bed is exposed to vacuum and vents the

adsorbed CO2 into space. Since this system also removes water from the ventilation loop, a condensing heat exchanger and water separator for humidity control is no longer needed, reducing system complexity and the amount of microbial growth seen in the ventilation loop. However, the RCA swing-bed scrubber operates at a slightly higher power than Metox and can only operate in a vacuum, meaning a diﬀerent system will eventually be needed for Martian EVAs [13, 16].

Figure 2-6: A schematic of the Rapid Cycle Amine dual swing-bed carbon dioxide removal system for implementation in the xEMU. One bed is connected to the ventilation loop, adsorbing CO2 and water vapor, while the other desorbs to vacuum. Note the piston that changes which bed is exposed to vacuum and vents the adsorbed CO2 into space. When a bed becomes saturated they switch providing continuous operation without requiring the unit to be removed and cleaned on station. Image from Reference [16].

2.4.3 Thermal Management

Heat in a spacesuit must be released at an adequate rate to keep the astronaut at a consistently safe and comfortable temperature. Heat during an EVA comes from three sources: human metabolism, suit electronics and the external environment. Metabolic and electronic heat are discussed here. Environmental heat is discussed in

39 Section 2.1. This section then details the technologies used in the current EMU to handle heat removal including the Liquid Cooling and Ventilation Garment (LCVG) and sublimator. The next section then discusses technologies that could improve the a spacesuit’s thermal management capabilities.

Human Metabolic Rates and Requirements

Human metabolic heat can ﬂuctuate dramatically depending on activity level. The metabolic heat from a single lunar EVA can vary from 70 to 730 W depending on activity, with extreme peaks of 880 W in emergency contingencies [7]. This heat can be broken down into two categories: sensible heat and latent heat. Sensible heat is transmitted through the skin to the environment via conduction and convection, and generally remains stable over varying activity levels. Latent heat refers to the heat rejected from perspiration and expired water vapor [17]. Latent heat increases rapidly with activity level (i.e. a person sweats more), as shown in Figure 2-7.

Figure 2-7: The human body rejects both sensible and latent heat. Latent heat rejection increases with activity level as perspiration and exhalation increase. The LCVG in the EMU changes this relationship by increasing sensible heat rejection and preventing sweat production even at increased exertion levels. Image from Reference [17].

The current LCVG and EMU alter this relationship by adjusting to keep skin temperature constant even at higher work rates so that sensible heat continues to be the dominate heat transfer mechanism. By actively drawing heat away from the skin, latent heat rejection is eliminated and sweating prevented. A system that relies

40 on the body’s natural transition to latent heat rejection at higher workloads may be more eﬃcient and reduce consumables but would also require more drinking water. It may also require a new mechanism for collecting and ﬁltering moisture condensate. This idea is explored in Chapter 4 of this thesis.

Electronics Requirements

The other heat source to be handled by a suit’s thermal management system is electronics. PLSS electronics heat will likely decrease with technological advancements, but generally runs at a steady 120 W [18]. The current requirement for NASA’s next EVA spacesuit, the xEMU, is 100 W with a goal of 75 W or less. In total, this means that a spacesuit must be able to reject between 190 and 850 W. The current EMU has a requirement to be able to reject approximately 250 W in steady-state performance for an 8-hour EVA [11]. The Apollo EMU requirement included about 270 W of cooling for an 8-hour EVA with a 19°C internal temperature maintenance, and a 250K heat sink temperature [12]. This represents a conservative baseline for future suit designs.

Liquid Cooling and Ventilation Garment

Figure 2-8: The Liquid Cooling and Ventilation Garment. Credit: NASA

41 Liquid cooling garments have been the accepted method of cooling EVA astronauts since the Apollo era. Both the EMU and the Russian Orlan suit utilize Liquid Cooling and Ventilation Garments (LCVGs). The LCVG, shown in Figure 2-8 is made of nylon and spandex interlaced with flexible rubber tubing for pumping cool water across an astronaut’s skin to absorb heat. Water enters the LCVG using the Liquid Transport Circuit through a manifold connector on the astronaut’s lower back. Water flows through the 91 meters of tubing at a nominal rate of 1.8 kg/min (4 lbs/min). This rate can be adjusted by the pump according to the metabolic thermal load detected. The water returns to the PLSS and is cooled via a heat exchanger in the sublimator before again circulating through the LCVG. The astronaut can reduce the cooling effect through the Display and Control Module (DCM) by bypassing the water flow from sublimator or even bypassing the LCVG itself [11].

Figure 2-9: Airﬂow is shown through the Liquid Cooling Ventilation Garment. Image from Reference [11].

42 The LCVG is also part of the Oxygen Ventilation Circuit reviewed in Section 2.4.1. As shown in Figure 2-9, inlet vents at the wrists and angles draw warm air back into the PLSS for scrubbing, cooling, dehumidifying and fresh oxygen injection [11].

Sublimation

Heat is rejected from the current EMU using a sublimator, a technology adopted from the Apollo A7 suit. Water from a feedwater circuit passes through the steel plates of the sublimator and are exposed to vacuum. One of the plates is porous and exposed to the vacuum of space. The water freezes through the pores, blocking the ﬂow, and sublimates into space. Thermal energy is removed as mass dissipation, leaving the plates at a cold temperature below the freezing point of water. Water from the separate Liquid Transport Circuit and oxygen from the ventilation circuit passes beneath the cold plates of the sublimator and are cooled before returning to the suit [11]. A schematic of the EMU sublimator is shown in Figure 2-10.

Figure 2-10: A cross section schematic of the EMU sublimator. Feedwater is frozen and sublimated through a porous steel plate into space, leaving the metal very cold. Water from the Liquid Transport Circuit passes through an aluminum housing beneath the sublimator and is cooled to about 13 °C (55°F). Underneath the feedwater, the oxygen in the ventilation ﬂow is cooled, condensing and separating moisture. Image from Reference [11].

The sublimator has a long flight heritage and is effective at rejecting large metabolic heat loads, but it also consumes a significant amount of water. A single 8-hour EMU

43 EVA burns though about 3.6 kg (1 gallon) of feedwater [19]. This mode of operation because less tenable for long duration missions without resupply. It will be replaced in the xEMU by the Spacesuit Water Membrane Evaporator (SWME), discussed in the next section [20, 13, 21].

2.4.4 New and Advanced Thermal Management Technolo- gies

As discussed, the LCVG and sublimator have a long succesful flight history. However the cooling system in NASA spacesuits has remained largely unchanged in the last 50 years. Clearly there exists an opportunity to leverage new technologies to reduce the mass and improve the efficiency of the thermal management system in future EVA suits. Improvements can be made in all life support subsystems including atmospheric revitalization and CO2 scrubbing, but special attention is paid in this thesis to thermal systems because some new cooling technologies could be implemented more effectively in a mechanical counterpressure suit like the BioSuitTM than in a gas pressurized suit like the EMU. A primary goal of these advanced thermal management technologies is reducing or eliminating the consumable mass needed to provide cooling. Discussed first in this section is the Spacesuit Water Membrane Evaporator, the system that will replace the sublimator in the xEMU. Additional proposals for eliminating consumable mass include the Spacesuit Evaporator Absorber Radiator (SEAR) and radiator-only based approaches first introduced in Hodgson’s Chameleon Suit architecture [22, 23]. A detailed exploration of advanced life support systems currently in development across industry and academia can also be found in Reference [24].

Spacesuit Water Membrane Evaporator (SWME)

The PLSS built into the xEMU will include several updates from the original Apollo and Shuttle era suits, including the integration of a Spacesuit Water Membrane Evap- orator (SWME) that replaces the sublimator. The SWME utilizes 27,000 thin-walled, hollow, 300 µm diameter polypropylene ﬁbers that provide about 1.1 m2 of open pore.

44 The small pores in the fibers prevent liquid from leaking out while presenting very low resistance to evaporation and vapor flow. The circulating water evaporates due to the low pressure on the outside of the SWME. The evaporation cools the remaining liquid as it returns to the LCVG. Separate liquid transport circuits for LCVG cooling and sublimation are no longer needed. However, this means all contaminants will remain in the cooling line, requiring a periodic scrubbing of the water circuit. The heat flux is controllable, and the SWME is pressure independent at start-up, operating over the whole range of suit pressures. The SWME can also degas the thermal loop water to prevent pressure build-up[13]. Testing began in 1999 at the Johnson Space Center, with the first full sheet demonstration in 2009. The Gen2 SWME was created in 2010 and has demonstrated a capacity of greater than 800 W of cooling capacity with over 200 hours of testing [21]. The SWME was tested on the ISS for the first time in 2019 as part of the Spacesuit Evaporation Rejection Flight Experiment (SERFE) [25]. The SWME will be more efficient and reliable than the sublimator it replaces. However, the SWME remains a mass-consumable life support system and can only operate in a vacuum, meaning a different system for heat rejection will be needed in future Martian EVA suits.

Spacesuit Evaporator Absorber Radiator (SEAR)

Another potential new heat rejection system is the Space Evaporator-Absorber-Radiator (SEAR), which utilizes a Lithium-Chloride Absorber Radiator (LCAR) to condense and absorb water vapor evaporated through the hollow ﬁbers of the SWME. The lithium-chloride solution is a strong desiccant that maintains a very low vapor pressure even at relatively high temperatures. The condensation reaction produces a lot of heat that can drive up the temperature of a connected exterior surface radiator panel that rejects heat into space. This is shown in Figure 2-11. The lithium chloride solution continues to absorb water vapor even at temperatures that are typically 30°C higher than the SWME temperature. This enables the LCAR to operate as a heat pump, rejecting heat at temperatures of about 50°C-60°C which signiﬁcantly reduces the size of the radiator required. Meanwhile the water

45 Figure 2-11: A schematic of how the lithium chloride cartridges would capture water vapor from the SWME, producing heat that can be rejected as radiation. Image from Reference [26].

itself remains captured in the LCAR system, meaning no mass is wasted. The LCAR designers expect the system to be able to reject 150 to 300 W depending on available radiator surface area and environmental conditions [22].

The concentration of lithium-chloride in the solution begins at about 95% and slowly dilutes to 45% over the course of an 8-hour EVA, at which point the LCAR ceases to be eﬀective. The absorbed water can be recovered after the EVA by baking the LCAR at 100°C-120°C, which is a high enough temperature to evaporate the water from solution [27]. This could be done using a regeneration oven inside the spacecraft or habitation module. A comparison of the independent SWME and SWME with LCAR systems is shown in Figure 2-12.

The LCAR system dramatically reduces water loss as compared with an SWME- only system, but it is not perfectly closed loop. Venting of certain trace non-condensable gases is necessary and results in some water loss, but much less than direct water vapor venting. Venting water may also be necessary to increase cooling capacity in hot external environments, extreme metabolic cases, and at times when the LCAR solution approaches 45% concentration and ceases to absorb water vapor. Thermal vacuum, impact, and pressure tests have veriﬁed the performance of LCAR in lab settings, but research has paused in the last several years from a lack of funding [27, 26].

46 Figure 2-12: Comparison of heat rejection using SWME alone and the SWME with LCAR (SEAR). The SEAR system rejects heat by absorbing water vapor and radiating the latent heat, as compared with the SWME independent system that directly vents the water vapor [27].

Radiators

As opposed to spacesuits, spacecraft and robotic explorers use radiators for heat rejection. Radiators reject residual heat via infrared (IR) radiation emitted by a surface exposed to a heat sink with a much lower temperature. These systems do not require any mass rejection. Radiators typically provide a steady level of cooling, making them suboptimal for adapting to the dynamic heat loads of a spacesuit application. Addi- tionally, surface area constraints on rigid structures such as the PLSS backpack limit the amount of cooling possible, as seen in the LCAR system [28, 26]. However, recent technological developments have reopened the possibility of using a radiator-based approach for EVA thermal control. Early work investigating the application of passive radiators for zero-consumable heat rejection led to the Chameleon Suit architecture, proposed in the early 2000s by Hodgson [23, 29]. With this approach, the thermal control system was envisioned to allow for variable conductance between the skin surface and the surface of the suit using diﬀerent lofting techniques. The surface of the suit allowed for further adjustment of the heat ﬂow using variable electrochromic radiators along with mechanical modulating of the radiators using MEMS (microelectromechanical system)

47 louvers. Technological advancements in the succeeding decades have made such a concept more viable, but it is still limited by an inability to remove heat when the sink temperature is close to or greater than the desired skin temperature (as can happend often on the Moon), and it still requires battery power to operate [30, 23].

Electrochromic materials may help the problem of environmental temperature variation by allowing for a variably-emissive surface. The optical properties of electrochromic films can be controlled electrically using an applied voltage [29, 31]. With this capability, shaded radiator surfaces facing cold heat sinks could be set to maximum emissivity whereas surfaces exposed to hot conditions in direct sunlight could be set to low emissivity and absorptivity values. New technology has also allowed these devices to be installed on flexible substrates such as Kapton tape or polymer film. When a small bias voltage is applied, the material undergoes an oxidation- reduction reaction, and the surfaces IR emission properties change. The layers of an electrochromic device (ECD) function similarly to the anode, cathode, and electrolyte in a battery [32].

Emissivity is measured on a scale of 0 to 1, where 0 indicates that no IR radiation is emitted (and no heat dissipated), while a 1 indicates that the device acts as an ideal black body and emits radiation at its theoretical limit. Technological demonstrations by Ashwin-Ushas Corporation and Eclipse Energy Systems have demonstrated variable emissive radiators with emissivity values that can be modulated from 0.19 to 0.90 [33, 34]. The energy emitted is governed by the Stefan-Boltzmann Equation:

4 4 Qrad = σ(Ts − Te ) (2.3)

Where σ is the Stefan-Boltzmann constant, is the emissivity, Ts is the surface temperature and Te is the environmental temperature of the environment. Much more research and development needs to be done before these devices may be implemented in a spacesuit design.

48 2.4.5 Summary

The sections above have reviewed the life support systems of historical EVA suits including the Gemini, Apollo, EMU and xEMU suits. Then a deeper dive into life support subsystems including atmospheric revitalization, carbon dioxide removal and thermoregulation was presented with the technologies used in the EMU reviewed as a case study. Finally, several advanced technologies for thermal regulation were discussed. The next section introduces mechanical counterpressure (MCP) spacesuits and past research conducted on their life support systems. MCP suits represent a fundamental architectural change from the gas pressurized suits reviewed so far. They oﬀer several potential advantages including increased mobility and decreased suit and life support mass, but also present many new engineering challenges. No MCP suit has ﬂown in space yet, but several prototype suits have been researched in the past or are currently in development like the MIT BioSuitTM.

2.5 Mechanical Counterpressure Spacesuits

All implemented spacesuit designs discussed so far have relied on gas pressurization within a sealed garment to protect astronauts from the depressurization of a crew capsule (IVA), or the direct vacuum of space during extravehicular activity (EVA). This approach has decades of flight heritage and has proven reliable. However, a suit with a full body self-contained atmosphere has several disadvantages. It adds mass and volume, is sensitive to small tears and punctures, and inhibits locomotion and general movement of the wearer. There is a very high metabolic energy cost of moving in a gas pressurized suit as an astronaut must do mechanical work to compress the gas within the suit every time a joint is bent. This is a concern in planetary EVAs as joint stiffness causes difficulty in traversing terrain or conducting surface operations. An example of this is the ”bunny hop” gait used by Apollo astronauts on the Moon to compensate for reduced lower body mobility [35, 36, 37]. Overuse strain injuries can also occur. Mechanical counterpressure (MCP) is an alternative method of providing pressur-

49 ization. The idea has been explored since the 1950s, most prominently in Webb and Anis’s Space Activity Suit (SAS), a NASA Langley funded project in the late 1960s and early 1970s [38]. More recently, research has been ongoing at the MIT Human Systems Lab since 2001 on a suit architecture led by Professor Dava Newman called the BioSuitTM. In an MCP suit, tight ﬁtting elastic garments provide direct contact pressure on the human skin. It is thought that this approach could improve astronaut mobility during EVAs, eliminate the risk of small tear and puncture-associated depressurization, simplify thermoregulation for the astronaut inside, and reduce overall life support and suit mass [35, 36]. The only example of a PLSS developed for an MCP suit is the Space Activity Suit. Up to this point, very little research has been done in the life support system for the BioSuitTM.

2.5.1 Space Activity Suit

Dr. Paul Webb and James Annis developed a series of elastic Space Activity Suit (SAS) prototypes between 1967 and 1971 under a research contract at NASA Langley [38]. They began with an elastic sleeve and glove before advancing to a full suit. By the end of the project they had assembled ten different garment assemblies and several helmet and breathing system iterations. The final iteration was designed to supply 3.3 psi (22.7 kPa) of contact pressure and allow for positive pressure breathing up to 3.9 psi (26.7 kPa). It included up to 10 different garment layers consisting of the following:

• Slip layer

• Torso pressurizing breathing bladder

• Helmet-bladder restraint garment—two full body garments to aid in controlling helmet rise while providing up to 0.48 psi (3.3 kPa) each to the arms and legs

• Arm balance layer—0.58 psi (4.0 kPa) added to the arms to balance counterpressure with the legs

50 • Full-body bobbinet—Two garments, each applying 0.29 psi (2.0 kPa) to the torso and 0.48-0.58 psi (3.3-4.0 kPa) to the limbs

• Third bobbinet—optional bobbinet layer to support breathing pressures above 3.3 psi (22.7 kPa)

• Girdles—Two powernet girdles, each with gradient pressure application on the upper thigh (0.68 psi/4.7 kPa) and lower torso (0.29 psi/2.0 kPa), shown to be needed when breathing pressure exceeded 2.7 psi (18.7 kPa).

Figure 2-13: The Space Activity Suit, including a compensatory torso breathing bladder on the left and the full donned suit with helmet on the right [38]

Figure 2-13 shows the torso breathing bladder layer and fully donned suit. The suit had a mass of 6.5 kg with an 18 kg breathing system. This is much less than the 91 kg fully loaded Apollo A7L suit or a 141 kg ISS EMU. In all, the suit took 45-60 minutes to don with the assistance of test associates and an average of 6 minutes to doﬀ. The suit was tested in an altitude chamber with pressures as low as 0.4 psi

51 (2.7 kPa) where a test subject completed a variety of physical activities with many physiological measures recorded. Figure 2-14 shows the test subject participating in tests including riding a stationary bicycle in the altitude chamber. The subject showed no extreme adverse eﬀects and exhibited an overall improvement in mobility as compared to the Apollo A7L, though the SAS was tested without a bulky Thermal Micrometeorite Garment (TMG) cover layer [39]. However, many of the altitude chamber tests were cut short due to test subject discomfort including swelling in some extremities, water loss and elevated heart rate and—in one trial–a breathing system hardware failure [39, 38].

Figure 2-14: A subject completes mobility exercises in an altitude chamber. Exercises included running on a treadmill, riding a stationary bike (right), climbing ladders and even crawling. The large white PLSS backpack can be seen in both images. The PLSS system was rapidly developed and failed several times creating an unsafe test environment [38] .

The life support system was comprised of a liquid oxygen tank, a soda lime (sodium hydroxide and calcium oxide) CO2 scrubber, and a gas reservoir all contained in a thermally insulating polyurethane foam within the pack. Note that it did not contain any thermoregulation hardware due to the ability of an MCP suit to allow for natural cooling. Figure 2-15 shows an original schematic of the SAS life support system. The breathing system was rapidly developed and failed multiple times, including causing one subject to collapse during an altitude chamber test. Still, it represents the only full PLSS built for an MCP spacesuit and therefore serves as a useful reference [39].

52 It was sized to support four hours of oxygen consumption. The pack had a mass of 17.9 kg (29.5 lbs) fully charged. This included 3.0 kg (6.6 lbs) of liquid oxygen and

1.8 kg (4.0 lbs) of soda lime CO2 absorbent. The inclusion of a liquid oxygen supply is a notable departure from other EVA suits. In this case, the liquid oxygen absorbed heat from the exothermic CO2 absorption process and was converted to a gaseous phase. Regulators maintained a gas pressure of about 63 psi (434 kPa) in the gas reservoir with a purge valve if oxygen boiloﬀ caused any pressure spikes. This was stepped down by additional regulators to 4.0-4.5 psi (27.6-31.0 kPa) for breathing. Flow was powered by a centrifugal fan tuned to provide 6.2 cfm (0.18 m3/min) of oxygen [38].

Figure 2-15: Schematic of the life support of the Space Activity Suit. It included an insulated liquid oxygen tank, cooling coils around a CO2 scrubber, pressure regulators, relief valves and a centrifugal fan to pump air into the helmet. Note the absence of any thermoregulation hardware or feedwater circuits. Image from Reference [38].

53 2.5.2 MIT BioSuitTM

The MIT BioSuitTM is an MCP spacesuit concept that has been in development since the early 2000s. The BioSuitTM system aims to use patterning along the lines of non-extension to apply constrictive force without inhibiting movement, and active shape memory polymers to ease donning and doffing. The lines of non-extension were first investigated by Arthur Iberall in the 1964 allow non-elastic pressure application without inhibiting mobility by taking advantage of contours on the skin that rotate but do not deform or stretch during joint movement [40]. This would theoretically alleviate some of the problems of the Space Activity Suit that used only elastic layers to provide pressure and was very difficult to don and doff.

Figure 2-16: BioSuitTM illustrations and mock-ups with lines of non-extension patterning. Center image credits to the author and left and right images to Dava New- man.

Laser scanning has been conducted to assess areas of skin deformation during joint ﬂexion. This information was then used to construct lower leg prototype sleeves tested in a low pressure leg chamber [41, 42, 43]. Similar analysis was also done using optical motion capture to identify lines of non-extension [44]. Shape memory alloy coil actuators have also been tested with elastic and inextensible fabrics as a potential

54 solution to make the suit easier to don and doﬀ by producing a garment capable of constricting on demand [45]. Recent work has shown that shape memory polymers may perform better than the shape memory alloys at actuating along these lines of non-extension to tension the suit after it is donned and relaxing prior to doﬃng [37]. As shown in Figure 2-16, mock-ups have been created that show the approximate patterning of these lines of non-extension.

Figure 2-17: Conceptual renderings of a conformal helmet and modular portable life support system for the BioSuitTM created by Michal Kracik.

Less research has been dedicated to the life support system for a BioSuitTM. Kracik et al. developed concepts for a conformal helmet and modular PLSS with components that could easily be replaced and consumables that could be swapped out during EVA [46]. One of these renderings is shown in Figure 2-17. The conformal helmet provides a more natural feel and provides more freedom of movement, but one of the biggest challenges is creating a sufficient pressure seal with the rest of the suit. Kracik sketched out a design with a rubber neck seal inspired by diving dry suits that lay underneath a helmet locking ring and provided an airflow connection to a torso breathing bladder. The helmet ring was required to have a latch and pin joint because it was too small to fit over the head and would have to be closed around an astronauts neck after donning.

55 Figure 2-18: A neck seal and breathing bladder integration concept developed for the BioSuitTM. The split helmet ring would open and close to allow it to ﬁt over the astronauts head during donning and doﬃng

Figure 2-18 shows a rendering of the neck seal and breathing bladder connection. This geometry served as inspiration for later work in the design of a compensatory breathing bladder prototype, presented in Chapter 3.

2.6 Summary

This chapter presented a review of the literature relating to spacesuit life support systems. We began with a review of the environmental conditions experienced in EVA locations including Low Earth Orbit, the lunar surface, and the Martian surface. Then we investigated the high level requirements that a spacesuit life support system must meet in order to keep an astronaut safe in these extreme environments. Next, we reviewed the history of planetary EVA suits and their respective life support systems including the Gemini G4C, the Apollo A7L, the Shuttle and ISS EMU, and the xEMU. Then we discussed each life support subsystem in more detail, including atmospheric maintenance and revitalization, carbon dioxide scrubbing, and thermal management. The past and current EMU technologies for these subsystems were presented. Several future technologies were also reviewed, including the Rapid

56 Cycle Amine (RCA) system for CO2 scrubbing, and the Spacesuit Water Membrane Evaporator (SWME), Spacesuit Evaporator Absorber Radiator (SEAR), and variably emmisive electrochromic radiator concepts for thermal mangement. Finally, we introduced mechanical counterpressure spacesuits including the Space Activity Suit (SAS) and the MIT BioSuitTM and discussed prior research into the life support systems of these suits. This chapter sets the context for the purpose of this thesis: to advance the development of a life support system for the BioSuitTM. Next, in Chapter 3, we present the design, development and initial testing of a compensatory breathing bladder for the BioSuitTM. This bladder serves a crucial role in an MCP suit by accommodating chest volume changes, easing the eﬀort of breathing, and equally distributing contact pressure from the suit across the torso. Then in Chapter 4, we investigate the thermoregulation system of the BioSuitTM by conducting thermal modeling of an MCP suit in lunar and Martian EVA conditions. Chapter 5 summarizes how both of these contributions advance and reﬁne crucial elements of a life support system for the BioSuitTM.

57 58 Chapter 3

A Breathing Compensation Bladder for the BioSuitTM

3.1 Background

Torso breathing bladders have a storied heritage in both high-altitude flight suits and spacesuits. Originally developed to protect against positive pressure breathing in fighter pilots in the 1940s and 1950s, the concept was adopted and applied in the 1970s for MCP spacesuits, beginning with NASA’s Space Activity Suit (SAS). The SAS bladder served to allow normal chest expansion during breathing. Most recently the BioSuitTM project at MIT has sought to revive MCP spacesuit research and development using modern materials and design processes. Part of this effort includes a new breathing bladder. While this counter-lung technology has a history, it has been several decades since an attempt has been made to develop the concept for a new suit. In this section we provide an overview of breathing bladder applications in aviation, the SAS, and the BioSuitTM. This overview is also summarized below in Table 3.1. Then this chapter presents the physiological and functional requirements of a breathing bladder in an MCP suit. Finally, this chapter details the design and fabrication process for a new prototype, and presents the setup and initial results of laboratory testing.

59 Table 3.1: Overview of Historical Applications of Volume Compensatory Breathing Bladders in Aerospace

Suit System Years Suit Type Bladder Implementation References Ernsting, 1966; Partial pressure Bladder used to decrease S-1, MC-1, MC-3 1940s-1950s Byrnes, 1999; ﬂight suits positive pressure breathing Jenkins, 2012 Prototype MCP Bladder used for breathing Space Activity Suit 1968-1971 Webb, 1971 spacesuit volume compensation Minimum of 20.7 kPa (3 psi) Conceptual MCP MIT BioSuit 2001-present absolute applied pressure Kracik, 2011 spacesuit to the body

3.1.1 Partial Pressure Flight Suits

Compensatory chest breathing bladders have been a facet of partial pressure and full pressure high altitude flight suits since the 1950s. They were first introduced to counteract the difficulty and danger of positive pressure breathing. Positive pressure breathing occurs when a pilot at altitude receives breathing gas from an air supply that is at a higher pressure than the surrounding atmosphere. In this scenario, the internal pressure in the lungs and bloodstream is higher than that felt on the outside of the chest and body. At altitudes above 13,000 meters (43,000 feet), a pilot must breath pure oxygen at positive pressure to get enough of the gas into the lungs to facilitate gas exchange in the alveoli. Breathing from a positive pressure gas supply means that a pilot has to forcefully exhale with each breath; a very tiring task. This is the case even with seemingly small differentials as little as 0.5 psi (3.4 kPa) [48]. As described by Byrnes in the book Blackbird Rising, “This pressure breathing is conducted under very low mask pressure, usually equivalent to the weight of a column of water about five inches high. Imagine blowing a five-inch plug of water out of your snorkel in the swimming pool each time you exhaled, and doing that for a couple of hours or more” [49]. In addition to breathing difficulty, there are other problems stemming from high intrapulmonary (internal lung) pressure including hyperventilation, overly distended and collapsed lungs, edema, blood pooling, and reduction in effective blood volume[50, 48]. To counteract this, several American aviation medicine labs began developing partial pressure suits in the 1940s, most notably the S-1, the first full body G-suit

60 Figure 3-1: The S-1 partial pressure suit developed at USC in the 1940s. Hollow tubes (’capstans’) along the legs and arms were inflated to tension the fabric and apply pressure to the pilot. Copyright David Clark Company [47]. with inflatable capstan tubes developed by Dr. James P. Henry at the University of Southern California. This suit is shown in Figure 3-1. The capstans would inflate when at altitude, increasing the tension in the suit fabric and applying more pressure to the pilot, minimizing the difference between the external and breathing pressures.

Inflatable chest bladders were developed around the same time. In 1942, Dr. Alvin Barach of Columbia University developed the first fully inflatable vest that extended the amount of time a pilot could breathe under positive pressure [47]. The bladder was connected to the breathing gas line and was inflated at the same pressure. As the pilot inhaled, the vest deflated, and when the pilot exhaled, the vest inflated, compressing the chest inward and helping the pilot exhale. Breathing, especially exhaling, remained hard work and represented a reversal of a human being’s normal respiratory physiology. In the mid to late 1950s, the David Clark Company developed

61 Figure 3-2: The MC-1 partial pressure suit developed by David Clark Company in 1955 included a chest and abdomen bladder, however it was determined to be too small to provide breathing beneﬁt and was replaced by the MC-3. Copyright David Clark Company and U.S. Air Force [47]. the MC-1 and MC-3 partial pressure suits which included an inﬂatable chest bladder. These suits (shown in Figures 3-2 and 3-3) allowed pilots to reach altitudes well over 30,500 meters (100,000 feet) for several hours at a time, but also blocked moisture ventilation from the torso and the pilots often reported that they became overheated [47].

Suit development then began to trend towards pressurized cabins with the only exception being unique cases like the U-2 or SR-71 Blackbird which still required full pressure suits to be worn. The S-100 suit built in the 1970s by the David Clark Com- pany for the U-2C aircraft included capstans and a coated nylon breathing bladder that wrapped completely around the torso and was integrated into the neck ring of the helmet. Despite the existence of these suits, pressure breathing for any prolonged length of time remains diﬃcult, even with specially designed systems that assist with

62 Figure 3-3: Pilot Francis Gary Powers wears the David Clark MC-3 suit, ﬁrst introduced in 1956. The MC-3 was much more comfortable and eﬀective than its predecessor with a complete torso bladder, but was phased out of US Air Force by 1964 in favor of fully pressurized suits and pressurized cabins. Copyright Lockheed Martin and David Clark Company [47]. exhalation [47].

3.1.2 The Space Activity Suit

Torso breathing bladders have been explored in several MCP spacesuit projects as well. The Space Activity Suit (SAS) developed by Dr. James Webb and Dr. Paul Annis at the NASA Langley Research Center in the late 1960s and early 1970s is to date the only full true MCP spacesuit that has been fabricated and tested [38]. The SAS included a compensatory bladder (”external counter-lung”), eight layers of MCP fabric, a bubble helmet, and a self-contained life support system. The SAS was reviewed in more detail in Chapter 2. The fully donned Space Activity Suit and compensatory chest bladder is shown in Figure 3-4. In spacesuit applications, the compensatory bladder is needed less for assistance in the event of positive pressure breathing but for accommodating the changing chest volume during respiration and equalizing pressure across the torso. Without a bladder, fabric across the chest would stretch and tighten during inhalation making it

63 Figure 3-4: The Space Activity Suit breathing bladder is shown at left. The fully donned MCP suit is shown at right [38]. diﬃcult to take a full breath, and then relax and loosen during exhalation and provide lower pressure than the rest of the suit [39]. In a well-designed MCP suit, the fabric will provide a constrictive pressure equivalent to the gas pressure within the helmet across the entire body so that an astronaut is not positive pressure breathing at all. Webb and Annis highlighted the need for a compensatory breathing bladder in this passage from their report:

“Early in the project subjects had real difficulty in breathing. This was thought to be due to the effect of as many as four layers of heavy powernet fabric over the torso breathing bladder. Shallow breathing with tidal volumes estimated to be 250-300 cc per breathing required little effort. Work requiring increases in respiratory tidal volumes in excess of 1 liter per breath could not be maintained for more than a few minutes. Tidal volume appeared to be limited near the high end of the vital capacity...Later

64 Figure 3-5: The Space Activity Suit helmet design including the integration of the chest bladder into a rubber neck seal. The helmet attached to a solid ring mounted on a breastplate on the subject’s shoulders. A soft air passage on the inside of the ring connected the bladder to the helmet [38].

in the ﬁrst phase tests, the amount of elastic counterpressure furnished to the torso was reduced. This allowed the bladder to do the job of volume compensation–that is, to expand with exhalation and contract with inhalation. When the breathing bladder supplied most of the torso counterpressure, the breathing was subjectively eﬀortless and respiratory rates were only slightly higher than those usually observed” [38].

The SAS breathing bladder had a maximum volume of 8 liters and a soft tubular passageway connecting it to the helmet airspace through the neck seal. The size of the

65 airway (20 cm2 when fully inﬂated) was intended to make pressure drops during the respiratory cycle negligible. It was made of rubberized neoprene-nylon cloth and sewn on top of a base comfort layer and underneath a non-stretch Nomex outer layer to restrict the bladder’s expansion so as to apply contact pressure to the wearer. Similar to the MC-3 suit, the nylon bladder was not permeable to moisture and the subject built up a considerable amount of latent heat around their chest during testing, but it otherwise functioned as intended. The helmet neck seal is shown in Figure 3-5, and the interconnecting airway between the helmet and the bladder is shown in Figure 3-6.

Figure 3-6: The interconnection of the Space Activity Suit torso pressurizing bladder and helmet shown in proﬁle view [38].

66 3.1.3 MIT BioSuitTM

Webb’s Space Activity Suit has provided important guidance in the development of the MIT BioSuitTM. Kracik et al. included a compensatory external chest bladder in a conceptual conﬁguration design of a helmet and life support system. A computer rendering of this BioSuitTM and the shape of the chest bladder is shown in Figure 3-7. In this design, the bladder would be pneumatically continuous with the air supply in the helmet, similar to Webb’s design. To address the problem of sweat build up, it was proposed that the bladder be cut out of a treated Gore-Tex fabric that is semi-permeable to water vapor while still remaining airtight [46]. If the bladder were permeable to perspiration it is possible that this would require additional humidity control measures in the helmet ventilation system as moisture could build up in the bladder and reach the helmet through the connecting air channel.

Figure 3-7: A conceptual rendering of the full BioSuitTM assembly. At right is a proposed chest bladder design with integrated neck ring that would lock into the helmet seal. Renderings by Michal Kracik [46].

Webb and Annis as well as Kracik’s work inspired the construction of the physical prototype presented in this thesis. We now explain the steps that have been taken

67 so far in the design, fabrication, and testing of a model breathing bladder for the BioSuitTM.

3.2 Design and Analysis

3.2.1 Physiological Requirements

The requirements for an MCP suit breathing bladder stem from issues related to pressure and volume. As discussed in Chapter 2, 3.5 psi (24.1 kPa) of pure oxygen is a safe minimum for breathing pressure in a spacesuit to enable adequate gas exchange in the lungs [10]. The contact pressure applied by the MCP suit should match the breathing pressure. This means the bladder should apply at least 3.5 psi (24.1 kPa) of pressure to the chest and also match the gas pressure in the breathing system. Deviations of up 1.0-1.2 psi (6.9-8.3 kPa) are acceptable for brief periods but can cause problems like positive pressure breathing after several minutes [39]. Finally, the contact pressure on the chest should spatially vary by less than 2.0 psi (13.8 kPa) with a goal of less than 0.25 psi (1.7 kPa) to prevent localized swelling [51]. For an EVA suit application, the breathing bladder volume must be sized to match an astronaut’s lung capacity. Work done to map the tidal change of the chest cavity has shown that volume changes vary quite signiﬁcantly depending on aerobic demand. Quiet breathing exchanges only about 0.5 liters of air. This increases to about 1.0 liters for fast, shallow breathing. For the average male, vital lung capacity (maximum inhale-exhale) is about 4.25 liters [52]. This is shown in Figure 3-8. Four liters was chosen as the design starting point because this captures all natural respiratory patterns including breathing during exercise. Note that this is 50% less than the bladder capacity of the Space Activity Suit. The reduced volume is appropriate given the actual requirements and the fact that the maximum SAS bladder volume was measured without the restraining layer of the suit in place. There is also a spatial distribution in volume changes across the torso. During quiet breathing, the pulmonary rib cage (above sternum) accounts for 29% of the

68 Figure 3-8: Average lung volume (VL) changes during quiet breathing (QB), metronome-paced tachypnea (MT) i.e., fast shallow breathing, and maximal forced inspiratory and expiratory manueuver (MFIE) for a sample of 15 males. The dashed and solid lines represent two different measurement systems, Chest Wall Optical Re- flectance (CW) and hot wire Spirometry (SP). C.V. refers to the Coefficient of Vari- ation between the measurement systems. Note that volume change for MFIE is over 4 liters for adult males, while for quiet breathing it is only about 0.5 liters [52].

volume change, the abdominal rib cage (below sternum) accounts for 24% and the abdomen itself accounts for 47%. This distribution reverses during maximum inspiratory and expiratory breathing. In this case, the pulmonary rib cage accounts for 41% of the volume change, the abdominal rib cage 31%, and the abdomen only 28% [52]. The absolute volume change is about eight times larger for this breathing pattern, meaning that pulmonary rib cage expansion should be prioritized in the geometric design of the garment. For this design, it was assumed that the large bladder volume and elastic material would accommodate any spatial diﬀerences in expansion. For future designs that seek to minimize bladder volume, this is an area that warrants further thinking. The bladder could be shaped so that when inﬂated, it is thicker around the upper rib cage, allowing for more chest displacement when the astronaut

69 inhales. Additionally, more elastic materials could be woven into the MCP layer at this location, giving the suit more stretch in the upper chest.

3.2.2 Design

The new BioSuitTM bladder designs were built using surfacing techniques in SolidWorksTM (Waltham, MA). A three-dimensional bladder shape was built that conforms to the chest shape of a 50th percentile male. Two concepts were developed, the ﬁrst includes a bladder that directly integrates into a neck seal for the helmet of the suit. The second includes an air hose that directly connects from a port in the shoulder of the bladder to the rear of the helmet. Figures 3-9 and 3-10 show renderings of these two designs.

Figure 3-9: A rendering of the bladder with direct neck seal integration. Notice at right how the fabric of the neck seal is clamped between layers of the helmet ring while a small passageway connects through to the bladder creating a pneumatically continuous system.

The air hose solution is easier to manufacture, but a hose that passes through multiple garment layers poses logistical challenges for donning and doﬃng a suit. Additionally, the hose could become caught or entangled during EVA activities if not restrained. Direct integration into the neck seal presents a simpler logistical solution

70 but makes the helmet neck seal more difficult to engineer, especially if the helmet ring is conforming and tight fitting around the neck (as shown in Figure 2-18) and requires a joint to open and close during donning and doffing, rather than simply being wide enough in diameter to be placed over the head. As discussed in the next section, the direct neck seal air passage idea was ultimately built for the first prototype.

Figure 3-10: A rendering of the bladder with an air hose (on the right shoulder) connecting the bladder to the helmet air space.

The initial CAD bladder design had a volume of approximately 3.25 liters. This was later increased to 4 liters to match the design goal. The cross-sectional area of the neck seal opening was about 14 cm2. The diameter of the duct in the air hose design was 2 cm (6.28 cm2), close to the proposed speciﬁcation for an alternative design of the Space Activity Suit ( 3 cm hose diameter) included in Webb’s report [38]. These three-dimensional designs were converted to two dimensional shapes for cutting, stitching and press-forming using ExactFlatTM (Toronto, ON), a CAD tool that optimizes 2D shapes from 3D surfaces [53]. The ExactFlatTM software builds a mesh over each piece of a three dimensional surface and then runs an algorithm to adjust the two dimensional shape to minimize the mesh strain given selected material

71 properties. These shapes are then converted to exportable files for CNC cutting machines. Figure 3-11 shows an example of the ExactFlatTM workflow. Pieces are selected from the surface then flattened and optimized. Next, edges from the 3D shape are matched in the 2D space and pieces are arranged for file export.

Figure 3-11: An example of the ExactFlat optimization process to produce two dimensional shapes that can be assembled into a three dimensional garment. The program starts by building a mesh on the surface (1). Then surface pieces are sepa- rated and flattened while minimizing strain in the mesh (2). Finally, the shapes and dimensions of the flattened pieces are exported in a file format used for cutting and manufacturing machines (3) [53].

3.2.3 Helmet and Breathing Bladder Air Flow Analysis

A first order analysis of the airflow inside of the helmet-bladder system was conducted using SolidWorks Flow SimulationTM tools [54]. Internal pressure inside a representative helmet volume was set at 4.3 psi (29.6 kPa) and air inflow and outflow were set at 6 cubic feet per minute matching the EMU operational target. Initial conditions of internal bladder pressure were set at

72 Figure 3-12: Internal ﬂow vectors show air entering the helmet at the rear, ﬂowing up and over the head and circulating into the bladder to equalize pressure before exiting through an outlet vent on the left side of the helmet.

1.5 psi (10.3 kPa) above and below the helmet pressure to represent bladder volume during inhalation and exhalation respectively. This represented a first order ideal gas approximation of pressure change given a control volume change of +/- 1 liter. Figure 3-12 shows the flow circulation in the model in the high pressure case. This simulation confirms that airflow exchange will occur even with a small interconnecting airway. Further, small pressure differentials between the bladder and the helmet will form during the breathing cycle but these will be less than 0.29 psi (2 kPa) and dissipate in seconds. This is acceptable by the recommendations made by Carr that spatial

73 pressure variations on the body be less than 0.25 psi (1.7 kPa) with up to 2.0 psi acceptable for short durations [51]. Figure 3-13 shows the initial pressure distributions in the helmet-bladder system at the point of maximum inhalation and exhalation. This is an imperfect simulation because it includes solid surfaces and does not model the changing volume of the bladder, but it can capture the initial conditions of the system and predict responses within a short time frame. These simulations may be refined in the future to improve and verify future design iterations. In actuality, the pressure differentials and flows between the helmet and bladder are likely to be much smaller because of the smaller volume change of normal breathing (0.5 liters) as compared to heavy breathing (>1 liter). This means pressure fluctuations in the helmet will be smaller than in this simulation (perhaps even negligible) and the airflow in the helmet will perform CO2 washout effectively.

Figure 3-13: Pressure contour maps for low bladder pressure (left) and high bladder pressure (right) show that helmet pressure can be maintained safely throughout the breathing cycle.

74 3.3 Fabrication

Fabrication of the bladder prototype was completed in collaboration with the D-Air LabTM of Vicenza, Italy [55][56]. An oﬀshoot of parent company Dainese, the D- Air Lab has extensive experience engineering personal airbag systems for motorcycle, MotoGP, and downhill skiing suits. These airbags are typically manufactured with an internal fabric laminated between two layers of thermoplastic polyurethane (TPU) or TPU-coated nylon. During the lamination process a valve is inserted for inﬂation and pressure measurement of the airbag.

Figure 3-14: Modeling of the breathing bladder geometry and neck seal at the Dainese D-Air LabTM. On the right is a cross-sectional model that shows how the inner surface will be joined to an airtight rubber seal around an astronaut’s neck while the outer surface will be attached to a rigid ring for mounting the helmet. Pictures used with permission of D-Air LabTM.

At the D-Air Lab, the team worked on a design that optimized ergonomic ﬁt while providing the required volume and surface coverage. We also developed a new connection mechanism between the helmet and the bladder and considered the best

75 materials for both fabrication and operation. It was decided that due to airflow considerations, it was best to design a bladder that had a direct passage connection from the helmet rather than a separate air hose. Geometry prototyping and the creation of the neck seal shape was first done on a mannequin model as shown in Figure 3-14. From these geometry prototypes, the initial two dimensional sketches produced in the earlier CAD modeling were modified to include accommodation for a neck seal that includes a tight fitting rubber sleeve around the neck and an outer bladder layer that is clamped into a rigid helmet ring.

Figure 3-15: BioSuitTM Breathing Bladder Prototype. The bladder is made of a semi-elastic and puncture resistant thermoplastic polyurethane (TPU). The interior layer is molded to a latex rubber neck seal normally used for dry suit diving. The outer layer is molded to a rigid ring onto which the helmet will be attached, forming a complete seal. The image at right shows the bladder inﬂated. In operation it would be constrained under an MCP layer causing the bladder volume to conform to the torso

The drawings were delivered to the airbag manufacturing division at Dainese. In consultation with the technicians there, it was decided to use only the TPU laminate material and not the TPU-coated nylon because the bladder would have a very low

76 operating pressure (<4.3 psi) compared to their typical airbags. Dainese normally installs internal microfilaments in their airbags that lay flat during deflation and maintain bladder shape during inflation. However, the breathing bladder requires flexibility rather than rigidity in its inflated state. Additionally, these microfilament fibers are a bio-hazard if inhaled so they were not used. The final product is shown in Figures 3-15 and 3-16. The interior TPU layer is molded to a latex rubber neck seal normally used for dry suit diving. The outer layer is molded to a rigid wooden ring onto which the helmet can be mounted, forming a complete seal and leaving an open air passage to the bladder. The remaining air volume around the neck will be compressed under the MCP garment. The fully inflated bladder has a volume of approximately 4 liters. Unrestrained, the bladder can stretch to a larger volume but in practice the volume will be constrained by the pressure garment worn over the bladder.

Figure 3-16: BioSuitTM Breathing Bladder Prototype. The internal latex neck seal can clearly be seen around the model’s neck at left. The spherical helmet that will be used for pressure testing is shown mounted on the mannequin in the center image. An astronaut would don the bladder ﬁrst before donning the MCP garment, creating an air volume to accommodate chest movement during respiration.

For testing, a simple transparent spherical polycarbonate bubble helmet was sealed onto the helmet ring with epoxy and silicone sealant and a rubber gasket was sand- wiched in between to keep the system airtight. Future prototypes may use more

77 durable materials, add edge tapering to better adhere to the human body, add seam lines to enhance ﬂexibility when an astronaut bends forward at the waist, or include a shoulder mount or harness for the helmet similar to the breastplate in the SAS. The design and fabrication of the breathing bladder prototype is also captured in detail in Reference [57].

3.4 Testing and Initial Results

Functional pressure testing of the breathing bladder was conducted using a compressed air supply, pressure regulator, and sealed helmet-bladder assembly as seen in Figure 3-17. The bladder is mounted on a plastic mannequin. Pressure sensors developed by the novel® Corporation (Munich, Germany) are placed on the chest of the mannequin underneath the bladder [58]. Then the torso is wrapped tightly in elastic rubber tubing. This method uses band tension to calculate circumferential pressure applied to a cylindrical body. This relationship is governed by the equation:

F = P br (3.1)

Where F is the tension force in the band, P is the applied pressure, b is the band width and r is the radius of curvature [42]. Applying this equation to a mannequin with a chest circumference of 90 cm (35 in), a band width of 5 cm (2 in), and a target pressure or 24 kPa (3.5 psi) yields a band tension of about 86 N (19.3 lbf). This method was attempted but ultimately abandoned because the mannequin’s torso had a cross section that too eccentric to be a practical approximation of a circle and pressure readings for a given band tension did not correlate well. The air compressor runs for the duration of the test, providing a steady air supply. A pressure regulator at the outlet of the compressor is set to 29.6 kPa (4.3 psi) to match the internal pressure of the EMU [11]. Next, a flowmeter reads the volume flow. The design goal was 6 acfm (170 liters per minute), again matching the EMU. The actual measured value during testing averaged 4 scfm (3.3 acfm) due to limitations in the pump capacity. Air flowed through a hose and entered through the

78 Figure 3-17: The bladder is wrapped underneath a layer of rubber tubing to apply compressive force to the chest. Sensors on the mannequin underneath the bladder measure the induced pressure. back of the helmet. Figure 3-17 shows a schematic of this test setup. The layering sequence of the test setup, with pressure sensors placed directly on the mannequin’s chest, then the bladder, then the rubber banding, is shown in Figure 3-18. After the bands were tension-wrapped, the airflow was turned on, the bladder inflated via the air passage through to the helmet, and the pressure sensors began recording data. Venting occurred through the backpressure relief valve. Figure 3-19 shows a timelapse of the inflation process.

After an initial test it was found that the contours of the mannequin made it diﬃcult to apply pressure uniformly. Some of the sensors showed promising readings while others showed zero values. For the next test, a single mat sensor with 256 cells from the same manufacturer was used to try and identify which areas of the chest gave the best pressure readings. The sensor was also placed on top of a ﬂexible plastic sheet that was wrapped around the mannequins chest to produce a smooth surface

79 Figure 3-18: Novel pliance® pressure sensors are placed on the chest of the mannequin (left). Then the bladder and helmet is donned (center). Finally, elastic bands apply pressure to the torso and the bladder is inﬂated (right).

with constant radius. This sensor and its placement on the mannequin can be seen in Figure 3-21. All pressure sensors were calibrated according to the manufacturer’s procedures prior to testing [58].

An additional problem encountered was helmet rise. When pressurized, the helmet tended to rise and press against the back of the mannequins head. The airspace in the bladder around the neck and shoulders also inflated and was difficult to con- strain. These two effects caused a contact pressure drop on the torso as a smaller volume of the bladder remained underneath the elastic bands. To solve this problem, a harness of yarn was tied around the neck ring and tightly knotted underneath the mannequin mount. This solution reduced the helmet rise observed but only represented a temporary fix. Future test setups will require a permanent solution perhaps including a restraining harness.

80 Figure 3-19: The bladder is underneath the black elastic pressure bands. The cables connect to pressure sensors on the chest of the mannequin. The air hose connects to the air compressor. Helmet rise during inﬂation is noticeable even with restraining yarn.

3.4.1 Individual Pressure Sensors Test

The individual sensors test (pictured in Figure 3-18) included 10 sensors placed on prominent and flat points on the the mannequins chest. During testing, nine of the ten sensor cells registered a signal and recorded data. Initial pressure readings (with only contact pressure from the elastic bands) ranged from 2.5 kPa to 17.5 kPa (0.4-2.5 psi). After the bladder was inflated, two of the sensors with a higher initial pressure recorded a decrease in pressure, but the remaining seven cells exhibited an increase in pressure. This is likely because the high pressure contact points were relieved by the bladder which more evenly distributed pressure. Nevertheless, the majority of the sensors failed to record pressure values above 8 kPa (2 psi), even after the system was fully inflated, meaning the system attained only about 50% of the targeted pressure value of 24 kPa (3.5 psi). The pressure response of the individual sensors in plotted in Figure 3-20. This data constitutes a pilot study and was not used for further analysis.

81 Figure 3-20: Initial test of pressure response using the 10 individual sensors. The average pressure reading increased after bladder inﬂation although points registering the highest values saw a decrease in pressure as it was distributed across the chest. This data is not rigorous and was not used for further analysis.

3.4.2 Mat Sensor Test

The novel pliance® mat sensor is comprised 256 sensor cells integrated as a single unit. Figure 3-21 depicts the sensor installed on the mannequin. Testing using this sensor exhibited a much higher pressure response, both in the number of cells recording and the maximum recorded values. Before the bladder was inflated, only 180 of the 256 cells (70%) recorded a pressure reading. After the bladder was inflated, this increased to 248 of 256 cells (97%). This increase in sensor response is visualized in Figure 3-22. The mean pressure detected by the full mat was 10.3 kPa (1.5 psi), but as with the individual sensors, specific cells recorded pressures reaching the target of 24 kPa (3.5 psi). Three cells reached this threshold and another six measured pressures within 20% of this value. While the peak pressure of several of the highest reading cells decreased after the bladder was inflated (the same phenomenon witnessed with the

82 single sensors), the average pressure for the whole sensor increased. A summary of key pressure values recorded is presented in Table 3.2.

Figure 3-21: The mat pressure sensor was placed over a plastic sheet to achieve a surface of constant curvature.

A plot of the the pressure reading of all individual cells from the mat sensor is shown in Figure 3-23. Clearly the response of each individual cell varied. Some cells recorded a small increase in pressure while others recorded a small decrease as pressure was distributed underneath the inﬂated bladder. The total average pressure and integrated force did increase after inﬂation.

83 Figure 3-22: A view of pressure cell readings before (left) and after (right) the breathing bladder is inﬂated. Some of the high pressure cells exhibit a small drop in pressure, but overall sensor activation increases as reﬂected in the higher number of colored cells at right. Cells in green and yellow meet the target pressure application. The three boxes at far right from top to bottom show the peak pressure, total applied force, and activated area of the sensor.

Figure 3-23: Mat sensor pressure response. The average pressure reading increased after bladder inﬂation although many cells recorded a small drop in pressure. The number of cells reading any amount of pressure also increased from 70% to 97%.

3.4.3 Discussion

A functional prototype of a compensatory torso breathing bladder for the BioSuitTM was fabricated and tested for the ﬁrst time. The bladder design was inspired by 84 Table 3.2: Statistical Summary of novel pliance® Pressure Data

Average Pressure Peak Pressure Standard Deviation Cells Activated Prior to Inﬂation 7.1 kPa 26.0 kPa 4.9 kPa 70% After Inﬂation 10.3 kPa 24.9 kPa 3.5 kPa 97%

Statistical summary of contact pressure data from the novel pliance® mat sensor before and after the bladder was inﬂated. Note that sensors reading zero were removed from the data set in computing these values.

prior work from Webb and Kracik [38, 46]. The bladder functioned appropriately and when inflated at a pressure consistent with current suit operational set points increased the average contact pressure applied to the chest from 7.3 kPa (1.1 psi) to 10.3 kPa (1.5 psi). However, there was a large variation between individual cells. This is reflected in the high standard deviation of recorded pressure values shown in Table 3.2. The failure to reach the goal of 24 kPa (3.5 psi) can be explained by a number of factors. The system did have several air leaks through the neck seal. It is an ongoing process to seal these leaks. There was also likely a pressure drop from the compressor through the air hose to the helmet inlet, so the air pressure inside the system may have been less than actually measured. The bladder was also sized using the anthropometric proportions of the author. The mannequin used for testing was proportionally too small for the bladder. Additionally, the hard contoured surface of the chest made it difficult to achieve a consistent contact surface for the pressure sensors to adhere to, resulting in inconsistent pressure readings and high variability between both individual sensors and cells of the mat sensor.

Nevertheless, the bladder successfully inﬂated in a system that included a human model, helmet and constricting mechanical pressure layer and produced an increase in pressure measured on the chest, proving the design concept is a viable method for achieving the contact pressure goal for a complete suit while allowing an astronaut to achieve appropriate lung volume without requiring unsustainable exertion. Addition- ally, several of the pressure cells within the pliance® sensor mat did measure contact pressures of 24 kPa (3.5 psi), indicating that with adjustments the system has the capability to match the functional pressure requirements needed in a full MCP suit.

85 Unfortunately this work was cut short and a sustained test campaign to collect data could not be conducted. Another aspect of data collection missing is the eﬀect of a changing chest volume on the experienced contact pressure. The primary purpose of the bladder is to provide a constant applied pressure while the chest expands and contracts. This functionality could not be tested. This means that the next step with this prototype design would be to simulate the cyclic chest volume change of breathing and measure the corresponding pressure on the chest to show that the system can maintain a constant contact pressure. This could be done with a mannequin that has an adjustable chest volume (a CPR mannequin for example), or a full human subject if full body compression can be achieved or a lower overall operating pressure used to prevent any interruptions to blood circulation that would occur if the head and torso are compressed more than the rest of the body.

86 Chapter 4

Thermal Modeling of BioSuitTM EVA on the Moon and Mars

4.1 Evaporative Cooling in an MCP Suit

A mechanical counterpressure suit may simplify the process of thermoregulation during EVA by eliminating the need for a Liquid Cooling and Ventilation Garment (LCVG). Instead of using water pumped through rubber tubing to draw latent heat away from the astronaut’s skin, the astronaut may be able to sweat directly through the layers of the suit similarly to how one may sweat through moisture-wicking ath- letic apparel worn on Earth. It is possible that this would reduce the amount of life

1 support system mass needed to support an EVA, perhaps even on the order of 3 [39]. Figure 4-1 shows the difference in garment layering between the EMU and a potential BioSuitTM. The BioSuitTM could consist of many fewer layers, reducing the mass of the suit. As discussed in Chapter 2, the current EMU relies on a Liquid Cooling and Venti- lation Garment to circulate cool water around the astronaut to draw heat away from the body. Additional ventilation tubes draw warm, moist air from the extremities for dehumidification, adding bulk to the suit. A feedwater circuit in the PLSS backpack freezes and evaporates water into space, removing heat from the water in the cooling line [11]. This method of cooling, using a sublimator, has a long flight history.

87 Figure 4-1: A layering comparison of the EMU (left) and a possible BioSuitTM layup (right). Note that the MCP suit does include the Liquid Cooling and Ventilation Garment (LCVG) which is the primary cooling mechanism in current spacesuits. The reduction in the number and density of suit layers in an MCP suit oﬀers potential mass savings.

However, a single eight-hour ISS EVA burns through about 3.6 kilograms of water (approximately 1 gallon), making water the single largest EVA expendable by mass [19]. For long missions to Mars lasting around 575 surface days, water consumed during EVAs could amount to 2,500 kg, or over 30% of the total ECLSS mass [22].

An MCP suit could circumvent the need for an LCVG and sublimator by allowing perspiration to evaporate directly into the surrounding atmosphere. Two types of water loss would occur: loss through passive skin diﬀusion and active perspiration. This idea was described in Webb and Anis’s report on the Space Activity Suit. They predicted an astronaut would lose approximately 360 mL of water through skin dif- fusion to a vacuum over the course of a 4 hour EVA (50 g/m2/hr for 1.8 m2 of body surface area). Based on prior physiological experiments, they noted that total sustained water loss rates should not exceed 2 liters per hour, but peak sweating rates of up to 4 liters per hour can be tolerated for short periods [38]. For each gram of water evaporated, 2.4 kilojoules of heat are removed. At a nominal water diﬀusion rate of 100 g/hr for human skin in a vacuum, 67 watts of heat will be removed, about half of resting metabolic heat. This is before sweating and conductive cooling are even introduced. At a sweat rate of 2 liters per hour, the body rejects heat through evaporation at a rate of 1350 watts, well above requirements. Vacuum or near vac-

88 uum environmental pressures would cause moisture to instantly vaporize, preventing sweat or ice build up on the suit surface. Physiological feedback and vasoconstriction would prevent overcooling [38]. The evaporative resistance of the clothing layers add another variable to actual water loss rates. During actual SAS human testing, water loss rates of 97 to 185 grams (0.097 to 0.185 liters) per hour were recorded, with higher rates experienced during low pressure altitude chamber testing. These data points suggest total water loss during an 8 hour EVA in near vacuum could amount to between 7 and 9 liters, necessitating the need for a reserve of drinking water in the suit. However, the data gathered was small in sample size, inconsistent between tests and therefore not likely reliable [39]. A summary of these parameters can be found in 4.1. Table 4.1: Comparison of metabolic cooling systems for the EMU, SAS and BioSuit

Number of Cooling Hardware Suit System Cooling System Water Loss Layers Mass LCVG and EMU 14 5.5 kg 3.6 kg per 8 hour EVA Sublimator SAS 8-10 Evaporative Perspiration 0 kg 7-9 kg per 8 hour EVA Evaporative Perspiration MIT BioSuitTM 3-10 Unknown kg ∼1.4-2 kg per 8 hour EVA and Radiators

The cooling hardware mass for the EMU refers to the wet LCVG (3.9 kg) and the sublimator (1.6 kg) [11]. Water loss rates for the SAS are projected based on small sample testing data and based on the simulation results presented in this chapter for the BioSuitTM.

It is also worth noting that the extreme radiation environments of space would still necessitate the need for a thermal micrometeoroid garment (TMG) to prevent the temperature of the MCP suit outer layer (and by extension the skin) from reaching extreme hot or cold values. The TMG wold also offer protection from dust on the Moon and Mars [39]. Discussion on sweating in an MCP suit has emerged in other suit proposals as well. Hodgson included water transport in several of his Chameleon Suit Architectures, noting the potential to remove a significant amount of PLSS mass. He also noted it may be possible to vary the evaporative resistance of clothing layers to adapt to different thermal environments [29]. Other discussion has focused on non-mass consumable methods including using suit surfaces as radiators to reject

89 heat [7, 28, 19]. These ideas are discussed in more detail in Chapter 2.

4.2 EVA Thermal Modeling

Historically, whole body human thermal modeling for space applications has been done using computer programs developed by Wissler, Hardy and Stolwijk in the late 1960s and early 1970s [59, 60, 61]. During the Apollo program, NASA relied primarily on the 41-node METMAN model developed by Kuznetz [59, 62]. For the the Shuttle and ISS EMU, NASA relied on an improved Wissler model with 225 nodes. In thermal modeling, the number of nodes represents the number of point solutions calculated and can be representative of model complexity. Incremental improvements for the last 40 years produced updated whole human models by Wissler and Fiala [59, 61, 63]. These models consisted of cylindrical elements representing body segments and considered countercurrent heat transfer between blood flows. Despite the history and volume of research built into these models, they still are run with Fortran code, a programming language not as commonly used today. For the work in this thesis, a modern industry software called TAIThermTM developed by ThermoAnalytics, Inc. (Calumet, MI) was used [64]. The goal was to evaluate the feasibility of using perspiration for thermoregulation in an MCP spacesuit by running several simulations of EVA in lunar and Martian environments with different garment configurations in the Human Modeling extension of TAIThermTM. TAIThermTM incorporates both the Fiala and Wissler model into its structure but provides a robust user interface with visualizations. This work built upon previous modeling completed by Vadhavkar, also using TAITherm. Vadhavkar simulated human heat loads during a four-hour Martian EVA for an astronaut wearing a simulated clothing layer and found that evaporative cooling is able to keep an astronaut comfortable up until the very end of the EVA [65]. This project aimed to refine and expand upon the modeling performed by Vadhavkar. An updated metabolic activity profile was simulated in the environmental conditions of both a Martian and lunar EVA.

90 Figure 4-2: An view of the TAITherm human model with diﬀerent body parts highlighted [64]

TAITherm has a robust built-in human model consisting of 28 body parts (shown in Figure 4-2), each of which has 16 distinct layers of bone, fat, muscle and skin tissue with unique thermal properties. The body parts include limb segments, and subsections of the head, torso and pelvis. The program’s inputs include human geometry; thermal, optical and evaporative properties of clothing; air temperature, wind speed, relative humidity, solar irradiation, activity level, and activity type. Clothing is added as additional layers on top of the human. Figure 4-3 shows a schematic of all the thermal processes included in TAITherm’s human model. TAITherm outputs the time varying temperature of each layer of the model as well as a comfort metric for each segment of the model based on a formula developed at the University of California Berkeley (referred to as the ”Berkeley Comfort Model”). This model is based on literature as well as human subject testing and uses uses skin temperature, temperature rate of change, mean skin temperature, and core temperature rate of change to assess overall comfort [66].

In total, the model built for this analysis consists of 5636 individual thermal elements used in the thermal ﬁnite element solver. The time step was set at 2 minutes,

91 Figure 4-3: An overview of all of the natural processes and elements included in the TAITherm Human Modeling program [64]. and the physiological results were saved at each time step. This small time step allowed for high resolution of thermal changes and also helped the model converge on a solution with fewer errors.

4.3 Using TAIThermTM

4.3.1 Human Model

The astronauts modeled are a 50th percentile male and a 50th percentile female. As described in the previous section, the human model consists of 16 layers of bone, muscle, fat and skin. Sweating and vasoconstriction were enabled in the model. The metabolic proﬁle used is the standard proﬁle used by NASA for qualifying elements of the PLSS for the next generation xEMU. It includes 20% margins on the high and low heat outputs expected during xEMU EVA [67]. A graph showing the activity

92 level over the course of an 8 hour EVA is shown in Figure 4-4. The activity profile used in this model was the first 240 minutes (4 hours) of the NASA model. This was done to both limit computation times and also because the entire range of human metabolic potential is encapsulated in the first half of the profile.

Figure 4-4: Human heat production profile used by NASA to define operational requirements for the next generation PLSS system [67]. The dashed line at 240 minutes indicates the ending point of the portion of this profile used in this simulation (0-240 minutes).

4.3.2 Spacesuit Model

Suit garments are added to the model as layers applied on selected surfaces with assigned material properties including thickness (mm), density (kg/m3), lateral con- ductivity (W/m-K), speciﬁc heat (J/kg-K), whole body evaporative resistance (m2- kPa/W), surface area augmentation (factor of surface area change from garment compared to bare skin) and whole body thermal resistance, also called thermal insulance (m2-K/W). Three diﬀerent clothing layups were simulated: a single nylon-spandex skinsuit layer, a skinsuit baselayer with a Boron Nitrite Nanotube (BNNT) infused jacket for radiation shielding, and the skinsuit with the BNNT and an Aerogel fabric layer for radiation and thermal protection. Boots were also worn in all simulations. The nylon-spandex skinsuit layer was based on prior research done by Holschuh

93 Table 4.2: MCP Suit Thermal Modeling Material Properties

Whole Body Whole Body Layer Thickness Density Speciﬁc Heat Material Thermal Resistance Evaporative Resistance [mm] [kg/m3] [J/kg-K] [m2-K/W] [m2-kPa/W] Nylon-Spandex 4.2 0.1 0.01 1150.0 1600 BNNT Fabric 2.0 0.25 0.05 8000.0 3500 Aerogel 1.5 0.5 0.01 900.0 0.002 Boots 5.0 0.22 0.08 1500.0 1600

Material property values used for diﬀerent suit layers in TAIThermTM Modeling [68, 69].

Figure 4-5: Conﬁguration of clothing layers used in the thermal simulations. Pink indicates where the nylon layer was applied. The green shows the BNNT over layer. The blue shoes where the boots were placed. The aerogel was also applied over the whole body the same as the pink skinsuit garment. and Kothakonda [70, 71, 37] as well as the material layers used in the Space Activity Suit. Webb and Annis used between 8 and 10 layers of fabrics named ”bobbinet” and ”powernet”. The bobbinet fabric was composed of cotton wrapped spandex ﬁber cores interlaced with nylon. The powernet fabric was composed of 75% nylon and 25% spandex [38]. Holschuh and tested a variety of MCP tourniquets using shape memory alloy coils to provide active tension to fabric samples. He found the best performing passive material for MCP among those tested was a fabric called ”jumbo spandex”. This material was composed of 0.7 mm thick layers of 90% nylon and 10% spandex. Holschuh recommended the skinsuit portion of a BioSuitTM be less than 5 mm in total thickness. Therefore he recommended between 3 and 7 layers of jumbo

94 spandex for the best performance [71]. Kothakonda is currently developing a single layer design with shape memory polymer (SMP) actuators to tension an inelastic fabric across the skin [37].

Figure 4-6: Test coupons of Boron Nitrite Nanutube (BNNT) and Aerogel radiation and thermal protection materials. On top is a BNNT sheet covered on both sides by a carbon nanotube (CNT) casing (for structural strength). On bottom is a combined BNNT Aerogel coupon encapsulated in polyethylene (a plastic with additional radiation shielding properties). The white side is the BNNT while the pale yellow side is the Aerogel. As evidenced by the image at bottom right, the product is ﬂexible, meaning it could be incorporated into a suit garment [69].

In general, the bulk thermal properties of the pressure layer will be dominated by the nylon-spandex fabric, even if there are embedded SMP actuators. Initial versions of the next MCP suit will likely require multiple pressure layers to achieve the required compression, but a single layer is the ultimate goal. Therefore, in this simulation, a single nylon-spandex layer 4.2 mm in thickness was added. A quick modeling comparison in TAIThermTM was conducted to justify this decision. Martian EVA simulations were run with a nylon-spandex base layer of 4.2 mm in total thickness

95 split into three layers, two layers, and a single layer. Core body temperature was not effected and mean skin temperatures varied by only 0.5°C after 4 hours, therefore the decision to use only a single layer was made. A graph of this result can be seen in Appendix B. Thermal results are very similar and a single layer decreases the computational time TAIThermTM needed to run simulations. The BNNT and Aerogel materials are the subject of current research for their potentially beneficial protective properties [69]. Figure 4-6 shows several test coupons of these materials that have been manufactured previously and may eventually be incorporated into protection garments. Figure 4-5 shows where the clothing layers were applied on the manikin. Radiative emissivity coefficients were set at 0.9, slightly less than the value of 0.98 for human skin but comparable with many white coatings used to aid in heat rejection in space applications [15]. The material properties of the three garment types are summarized in Table 4.2.

4.3.3 Weather Model

TAIThermTM is able to simulate a complex weather environment if a user uploads the requisite data in the proper ﬁle format. Weather models in the program are meant for Earth conditions, but by changing the temperature, pressure, humidity, and radiation values, it is possible to model the surface of Mars or the Moon. Table 4.3 shows a high level comparison of the weather conditions modeled for each planetary body and Figure 4-7 shows images of the surfaces of the Moon and Mars for visual reference.

Moon Weather

The lunar weather model was developed based on the expected average environmental conditions at the lunar South Pole, specifically Shackleton Crater, located at 89.8°S, 0.0°E. A South Pole landing site is specifically referenced in NASA’s Plan for Sus- tained Lunar Exploration and Development [2]. This is a significant departure from Apollo program landing sites, all of which were located at near-equatorial latitudes. The Apollo EVAs were timed to occur during lunar mornings [72]. Astronauts avoided

96 Table 4.3: Planetary EVA Weather Comparison

Average Atmospheric Relative Solar Diﬀuse Solar Solar Temperature Pressure Humidity Irradiance Component Position 0.3°alt. Moon -73°C 0.0 kPa 0.0% 1368 W/m2 5.2 W/m2 45.0°az. 86.0°alt. Mars -15°C 0.65 kPa 0.01% 504 W/m2 98 W/m2 10.0°az. 90.0°alt. Earth 14°C 101.3 kPa 20-100% 1000 W/m2 1000 W/m2 0.0°az.

A comparison of the average weather characteristics between the Moon and Mars used in the thermal modeling simulations. Average conditions for direct overhead sunlight at sea level on Earth are also included for reference.

the high Sun elevation angles and the full heat of solar noon where surface temperatures can reach 400 K (127 °C) and the cold of lunar night when temperatures can fall to 100 K (-173 °C) [73]. This was acceptable for short duration missions but a challenge for a permanent lunar base.

Figure 4-7: Images of the lunar surface (left) and the Martian surface (right). The lunar surface image was captured by the Apollo 17 crew in 1972 in the Taurus- Littrow Valley (20.2°N 30.8°E). No surface pictures exist of the South Pole region. The Martian surface image was captured by the Mars Science Laboratory Curiosity Rover in 2013 in Gale Crater (5.4°S 137.8°E). Image credits to NASA.

At polar latitudes, there is near constant sunlight for solar power production, relatively consistent surface temperatures, continuous direct to Earth communications capabilities, and potential access to water ice permanently frozen in shadowed craters

97 [74]. These factors make the lunar south pole an intended target for future human landings. The average illuminated surface temperature of the Moon in polar regions is about 200 K (-73°C) [73]. Direct solar irradiance on the Moon is approximately 1368 W/m2. Diffuse radiation is calculated as only 2 W/m2 using a lunar albedo (α) value of 0.08 and a solar zenith angle of 89.8°. The infrared heat flux from the lunar surface is about 5.2 W/m2 at the south pole [75]. Because of the Moon’s slow rotation rate and the the extreme latitude, these values were held constant throughout the duration of the EVA. The direct radiation is the same as during Apollo EVAs, but the diffuse component is much less, creating slightly more hospitable conditions.

Mars Weather

Figure 4-8: Atmospheric temperature and solar radiation levels for the simulated Martian EVA. Time prior to t=0 indicates steady state temperature conditions in the habitat prior to exit for EVA. Temperature and direct solar irradiance increase as the sun rises to a higher elevation over several hours. Diﬀuse solar radiation remains approximately constant. Image taken from Reference [65].

The Martian weather model was adopted from prior work by Vadhavkar, simu- lating a hot summer day in the southern hemisphere (-25°S) based on data from the Mars Global Surveyor and several Mars landers [65]. The air temperature and solar

98 irradiance for the four hour duration is shown in Figure 4-8. Temperatures range between -23°C and -6 °C and atmospheric pressure is only 0.65 kPa. Solar irradiance values increase as the Sun reaches a higher elevation in the sky. This weather pattern is meant to simulate a ”hot” EVA scenario during summer in the southern hemisphere. The latitude was chosen to maximize insolation, with 25 degrees corresponding to the axial tilt of Mars [65]. These conditions were chosen to simulate conditions that might be expected at a future Mars station built at a location with a climate warmer (and therefore more hospitable) than the Martian average.

4.4 Results

As discussed in the sections above, three separate simulations were run for each planetary environment: one with the astronaut wearing only a nylon skinsuit, one with a BNNT jacket on top of the skinsuit, and one with a skinsuit, an Aerogel insulation layer and a BNNT jacket. These test cases were run ﬁrst with a male manikin and then a female manikin. The results recorded and analyzed included the time-varying skin surface temperature and corresponding thermal comfort sensation provided by the Berkeley Comfort Model. These results are also visualized on the human model. Area averaged skin temperature and body core temperature were also recorded, as well as the sweat rate in grams per minute and the amount of time spend in any state of shivering. Table 4.4 shows the test matrix of simulations that were run. Table 4.4: TAIThermTM Modeling Testing Matrix

Clothing Data Collected Moon Nylon Skinsuit Skinsuit & BNNT Skinsuit, BNNT & Aerogel Skin and core temperatures, Mars Nylon Skinsuit Skinsuit & BNNT Skinsuit, BNNT & Aerogel Berkeley Comfort Model thermal sensation, heat ﬂuxes, and sweat rates

Data collected for the TAIThermTM EVA thermal modeling included different suit temperatures and heat outputs for different garment configurations. All tests were run with both male and female manikin models.

In normal conditions, the temperature of the human skin is between 33.5°C and

99 36.9°C [76]. The body’s core temperature is between 36.5°C and 37.5°C [77]. Hy- pothermia occurs when the body’s core temperature falls below 35°C [78], while hyperthermia occurs when the body’s core temperature exceeds 38.3°C [79]. These are the values used to assess safe conditions. These values are summarized in Table 4.5

Table 4.5: Core and Skin Temperatures for a Range of Human Thermal States

Thermal State Temperature Range Core: 36.5 -37.5 C Healthy ° ° Skin: 33.5°-36.9°C Hyperthermia Core: >38.3°C Hypothermia Core: <35°C Cold Numbness Skin: 20°-28°C Cold Pain Skin: 10°-20°C Frostbite Skin: <0°C

Core and skin temperature ranges for a variety of thermal states. The goal is to maintain astronaut core and skin temperatures in the Healthy range [76, 77, 78, 79, 80].

Additionally, heat ﬂuxes were recorded. TAIThermTM records the time varying heat ﬂux from the following sources [64]:

• Qm (Metabolism): Total rate of energy generated by the virtual manikin in response to the environment and activity level.

• Qconv (Convection): Rate of heat transfer to the virtual manikin’s skin and clothing by air, wind, water or other ﬂuids included as boundary conditions.

• Qrad (Radiation): Rate of heat transfer to the virtual manikin’s skin and clothing by infrared radiation exchange with the surrounding environment. It does not include solar loading, which is reported separately.

• Qsolar: (Solar): Rate of heat gained by the virtual manikin’s skin and clothing through direct, diﬀuse and reﬂected solar radiation.

• Qevap: (Evaporation): Rate of heat transfer occurring through latent heat of vaporization of perspiration.

100 • Qresp: (Respiration): Rate of heat transfer in the process of breathing. It includes both convective heat transfer and latent heat of vaporization when the breath increases the humidity of the exhaled air.

In the model created, the manikin also wore a pair of boots that was consistent between all tests. An enclosed climate-controlled helmet on the head could not be modeled. It was assumed that a helmet ventilation system (not modeled) would keep the astronaut’s head comfortable during any of these EVA scenarios. The head and face were also modeled as being covered with fabric layers so as to match the rest of the body. If this has not been done, the head would have been extremely cold in all simulations and skewed overall results. However, this means that thermal results speciﬁcally for the head should not be the focus of attention.

4.4.1 Moon Modeling

The extreme radiation environment of the lunar surface dominated the results of the lunar EVA thermal model. When only the nylon-spandex base layer was worn, body surfaces facing the sun experienced skin temperatures exceeding 60°C while surfaces in shadow dropped to below -40°C. The result is more severe than would be actually experienced because the astronaut does not move, meaning the sunlit and shaded halves of the body remain constant throughout the EVA. An actual astronaut would constantly be changing their orientation relative to the Sun while conducting EVA tasks, varying the exposure of different surfaces to the Sun. Nevertheless, the thermal load remains unsafely high. The heat loading is most extreme in the case of the astronaut wearing only the nylon skinsuit, shown in Figure 4-9. This figure shows a timelapse of the skin surface temperature and the thermal sensation experienced by the astronaut wearing only the skinsuit sampled at 0, 30, 60, 120 and 240 minutes. The astronaut is initialized with a clothing temperature of 20°C, representative of a person exiting a habitat kept at a comfortable 20°C temperature. The addition of the BNNT and aerogel protective layers began to add some shielding effects. The temperature visualizations for these tests can be found in Appendix

101 Figure 4-9: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Lunar EVA wearing only a nylon-spandex skinsuit. The extreme solar loading is evidence by the ”hot” side of the astronaut experiencing skin temperatures of up to 60°C while the ”cold” side temperatures drop to 12°C. These eﬀects would be more equalized in an EVA in which the astronaut does not maintain a constant orientation relative to the Sun. Clearly a radiation shielding garment is needed for direct solar exposure.

A. In all tests, the astronaut’s core body temperature increases but remains below the point of hyperthermia. The mean skin temperature and core temperature for each trial is shown in Figure 4-10. The addition of the BNNT jacket kept the torso at a lower and more comfortable average temperature. This suggests it successfully provides a level of radiation protection to vital internal organs. Skin temperatures remain the closest to normal with the full Aerogel insulation layer. Interestingly, the manikin began sweating to dissipate heat during both the skinsuit and the skinsuit with BNNT tests, but did not

102 Figure 4-10: A comparison of mean skin temperature (MST) and core body temperature for a Lunar EVA wearing three layering combinations: a nylon skinsuit only, a skinsuit with a BNNT radiation jacket, and a skinsuit with an Aerogel insulation layer and the BNNT jacket. sweat during the Aerogel clothing test. Even with higher mean skin temperatures, the Berkeley Comfort model also shows that while wearing Aerogel, the astronaut remains slightly cool, especially in the torso for the duration of the EVA. In contrast, while wearing only the skinsuit, the chest becomes extremely hot. This diﬀerence is likely caused by the ability of the insulation layer to prevent extreme hot and cold patches from developing on the body and allowing a more uniform temperature and thermal sensation. In the extreme case, once the body is insulated from the intense solar radiation environment, there is very little incoming heat ﬂux and a very low ambient temperature, meaning that the challenge in this situation becomes keeping the astronaut warm.

The heat ﬂuxes and sweat rates for the lunar EVA wearing the full Aerogel suit assembly are shown in Figure 4-11. This image shows that the two heat inputs are metabolic heat (modeled by the activity proﬁle) and solar heat (modeled by the lunar weather conditions described in the sections above). Heat is dissipated via

103 infrared radiation, atmospheric convection, respiratory exhaust and evaporation from perspiration. The sweat rate reached a maximum of 14 grams per minute while wearing only the skinsuit, equivalent to 0.84 liters per hour, falling within the safe sweat rate limit of about 2 liters per hour [38]. Total water loss due to perspiration for the four hour EVA was 0.891 liters. This water loss could easily be compensated by an in-suit drinking bag. Even for a full eight hour EVA, a total sweated water loss of 2 liters would be less than the approximately 3.6 liters used per EVA on the ISS now [19].

Figure 4-11: Heat ﬂuxes and sweat rate for a Lunar EVA wearing the full garment assembly (nylon skinsuit, BNNT and Aerogel). The manikin begins to sweat after portions of high activity depending on the amount of clothing insulation. Note that convective heat loss is zero because the Moon does not have an atmosphere.

The radiative heat dissipation is very large in this simulation, remaining at about 850 watts. This is certainly too large a value as the program also shows zero evaporative cooling despite a positive sweat rate that increases throughout the EVA. This is likely a program error stemming from the fact that convective heat transfer was

104 set to zero. Radiative heat dissipation would also be limited by available surface area in a real suit. The model does not account for a PLSS backpack or helmet that would reduce the amount of heat emitted. Nevertheless, these results support the hypothesis that evaporative cooling can provide a large portion of the heat rejection needed for a lunar EVA. Though these particular clothing conﬁgurations resulted in uncomfortable thermal conditions for the astronaut, they were survivable. It is clear that a full MCP suit for lunar EVA would need a thermal micrometeoroid garment to block harmful solar radiation while allowing perspiration to evaporate. This system would keep the astronaut a comfortable temperature while still allowing heat to escape in case of high stress activity.

Male and Female EVA Comparison for the Moon

A female astronaut experiences lower skin temperatures than a male counterpart for each suit conﬁguration. This results in increased cold discomfort as compared to the male astronaut. Figure 4-12 shows the mean skin temperatures and core body temperatures for all garment conﬁgurations for both the male and female astronaut overlaid.

The female manikin produced less metabolic heat and less sweat than the male astronaut, as evidenced in Figure 4-13. The female manikin reached a maximum sweat rate of 5 grams per minute, much less than the male manikin’s maximum rate of 14 grams per minute. This sweat rate falls within safe limits and only 0.369 liters of water were lost over the 4 hour EVA. As in the male manikin simulation, TAIThermTM recorded zero heat dissipation from evaporation despite the existence of perspiration. This is believed to be a programming error caused by convective heat transfer being set at zero to reﬂect lunar atmospheric conditions. Still, the overall heat rejection requirements for a female astronaut are less than a male due to a lower metabolic heat rate. This means that supplemental cooling systems for a female suit such as radiators could be downsized as compared to male suits.

105 Figure 4-12: Skin and Core Temperature Comparison for Male and Female Astronauts During Lunar EVA. Blue lines indicate simulations with a male manikin, red lines indicate simulations with a female manikin. Solid lines show results wearing only the pressure base layer, dashed lines the BNNT jacket, and dotted lines the Aerogel layer. Clearly a female astronaut exhibits lower mean skin temperatures throughout the EVA. For both male and female astronauts, the core body temperature remains within safe limits meaning that the EVA is survivable.

4.4.2 Mars Modeling

It was found that all three garment conﬁgurations created survivable thermal conditions for an astronaut in a simulated Martian EVA. However, only when the Aerogel insulation was added did the astronaut avoid becoming uncomfortably or even dan- gerously cold. In the ﬁrst simulation, with an astronaut wearing only a nylon-spandex skinsuit and boots, skin temperatures generally remained around 27°C, about 7°C below nominal. Core temperature remained at about 37°C, meaning the astronaut did not become hypothermic. Figure 4-14 shows the the skin surface temperature and the thermal sensation experienced by the astronaut while wearing only the nylon-spandex skinsuit over time.

106 Figure 4-13: Heat ﬂuxes and sweat rates for a female astronaut wearing the full Aerogel suit layup during the simulated Lunar EVA

Figure 4-15 shows the mean skin temperatures and core temperatures compared for all suit conﬁgurations. Skin temperature quickly drops in the cold Martian atmosphere before increasing slightly during the phases of the EVA with a high metabolic load. This is also shown in the Berkeley Comfort Model thermal sensation. The astronaut is very cold for most of the four hour EVA, and the model reporting that some form of shivering occurred about 90% of the time. When a BNNT jacket layer is added to the suit, the astronaut experiences a very similar thermal load. The torso is kept slightly warmer, especially during the high activity portions of the EVA, but overall the astronaut remains very cold. The human visualization for the EVA wearing the BNNT jacket can be seen in Appendix A. Localized hot spots on the torso are similar to what may be experienced with a breathing bladder in place. An air-ﬁlled bladder was not simulated due to software limitations, but it could cause uncomfortable perspiration build up if not designed correctly. Overall though, these results show that the radiation shielding jacket can

107 Figure 4-14: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Martian EVA conducted by a male astronaut wearing only a skinsuit. The astronaut experiences increasingly cold discomfort, oﬀset only a little by the metabolically strenuous phases of the EVA

be added to the suit assembly without signiﬁcantly aﬀecting the thermal properties of the suit.

Finally, with the addition of an Aerogel insulation layer to the suit, the astronaut exhibited much more stable skin temperatures and comfort. As seen in Figure 4- 15, both the mean skin temperature and core temperature where higher than the previous tests. After an initial drop due to cold exposure, skin temperature rose close to normal values, even surpassing 35°C at the end of the EVA after more stringent activity. The astronaut was initially slightly cold but became increasingly comfortable after an increase in metabolic heat during strenuous portions of the EVA. Figure 4-16

108 Figure 4-15: A comparison of mean skin temperature (MSK) and core body temperature for a Martian EVA wearing three layering combinations: a nylon skinsuit only, a skinsuit with a BNNT radiation jacket, and a skinsuit with an aerogel insulation layer and the BNNT jacket

shows the time-lapse of the skin temperature and thermal comfort sensation for the Martian EVA wearing the full suit assembly with aerogel insulation. The amount of time spend shivering drops to only about 15% of the EVA, all of which occurs in the early minutes of the EVA before any intense activity.

A full decomposition of the heat sources in the EVA is shown in Figure 4-17. Again, this image shows that the two heat inputs are metabolic heat (modeled by the activity proﬁle) and solar heat (modeled by the Martian weather conditions described in the sections above). Heat is dissipated via infrared radiation, atmospheric convection, respiratory exhaust and evaporation from perspiration. What sets the Aerogel simulation apart is the increase in evaporative cooling that occurs after high metabolic heat loading. This was also the only suit setup in which the modeled human began to sweat signiﬁcantly.

109 Figure 4-16: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Martian EVA wearing a skinsuit, BNNT jacket and Aerogel insulation layer. The astronaut experiences slight cold discomfort, which gradually dissipates after the physically strenuous portions of the EVA. Compared to the skinsuit only EVA, the astronaut remains much warmer and more comfortable.

The sweat rate over time is also shown in Figure 4-17. By the end of the EVA, the astronaut reaches a sweat rate of 15 grams per minute and is rejecting over 250 watts through moisture-wicking through the garment and evaporative cooling. This rate, equivalent to 0.96 liters per hour is within the recommended safe limit of 2 liters per hour by Webb et al. The sweat rate was integrated to produce a total sweat volume of 0.691 L for the 4 hour EVA. Again, this water loss could easily be compensated by an in-suit drinking bag. While the environmental conditions are diﬀerent on Mars than in Low Earth Orbit, the potential for water mass reductions and the elimination of bulky cooling loops from the suit make this a deﬁnite improvement.

110 Figure 4-17: Heat ﬂuxes and sweat rate for a Martian EVA wearing the full garment assembly (nylon-spandex skinsuit, BNNT and Aerogel). Sweating only occured when the manikin wore the full spacesuit assembly with Aerogel.

Male and Female EVA Comparison for Mars

As with the results from the lunar EVA modeling, a female astronaut experiences lower skin temperatures than a male counterpart for each suit conﬁguration. This results in increased cold discomfort as compared to the male astronaut. Figure 4- 18 shows the mean skin temperatures and core body temperatures for all garment conﬁgurations for both the male and female astronaut overlaid. The female manikin produced less metabolic heat than the male astronaut, as evidenced in Figure 4-19. The maximum energy dissipation rate from evaporation is just under 200 watts, over 50 watts less than the male astronaut. This makes sense given the smaller sweat rate. The female manikin reached a maximum sweat rate of 5 grams per minute, much less than the male manikin’s maximum rate of 15 grams per minute. This sweat rate falls within safe limits and only 0.238 liters of water were lost over the 4 hour EVA. The female manikin also spent slightly more time

111 Figure 4-18: Skin and Core Temperature Comparison for Male and Female Astronauts During Martian EVA. Blue lines indicate simulations with a male manikin, red lines indicate simulations with a female manikin. Solid lines show results wearing only the pressure base layer, dashed lines the BNNT jacket, and dotted lines the Aerogel layer. Clearly a female astronaut exhibits lower mean skin temperatures throughout the EVA. For both male and female astronauts, the core body temperature remains within safe limits meaning that the EVA is survivable. shivering than the male manikin in all suit conﬁgurations, illustrating the need for supplemental insulation and radiation protection in a female suit.

4.5 Discussion

The goal of this thermal analysis was to determine whether clothing representative of a mechanical counterpressure spacesuit such as the BioSuitTM could provide adequate thermal regulation in the harsh environments of the lunar and Martian surfaces. Table 4.6 summarizes key thermal data from the simulations. The simulation results support the claim that with adequate material properties, an MCP suit can rely on natural cooling processes to help keep an astronaut safe and thermally comfortable

112 Figure 4-19: Heat ﬂuxes and sweat rates for a female astronaut wearing the full Aerogel suit layup during the simulated Martian EVA for the duration of an EVA. Perspiration provided up to 250 watts of cooling during a Martian EVA, all at sweat rates well below the established healthy limit. On Mars, convective heat loss to the thin atmosphere also provided close to 300 watts of cooling. Convection is not available on the Moon however. For a suit that can handle up to 800 watts like the current EMU, additional means of heat rejection are needed, enough to provide up to 550 watts in addition to approximately 250 W available through sweating. In these simulations, most of the remaining cooling came from radiation, the other passive means of heat rejection. In reality, the ability of the suit to act as a radiator will be limited by available surface area as well as contamination from dust build up. Additionally, respiratory cooling comprised about 5-20% of heat dissipation for the EVAs. In an actual suit, this heat will need to be drawn out of the oxygen ventilation circuit in the helmet and PLSS backpack and rejected by another means. The need for up to 550 additional watts in heat rejection can likely be mostly accomplished using radiating surfaces directly on the body (as in these simulations)

113 Table 4.6: Summary of TAIThermTM EVA Modeling Results

Moon Ending Ending Max Time Max Total Suit Assembly Skin Temp Core Temp Evaporation Shivering Sweat Rate Sweat Loss Nylon-Spandex 30.64 C 37.54 C 0 W* 0.0% 13.75 g/min 0.891 L Base Layer BNNT & Male 31.57 C 37.23 C 0 W* 0.36% 5.94 g/min 0.226 L Base Layer Aerogel, BNNT 32.84 C 36.88 C 0 W 0.45% 0.0 g/min 0.0 L & Base Layer Nylon-Spandex 28.30 C 37.81 C 0 W* 0.0% 5.41 g/min 0.369 L Base Layer BNNT & Female 30.40 C 37.44 C 0 W* 0.33% 2.75 g/min 0.121 L Base Layer Aerogel, BNNT 31.68 C 37.03 C 0 W 0.41% 0.0 g/min 0.0 L & Base Layer Mars Ending Ending Max Time Max Total Suit Assembly Skin Temp Core Temp Evaporation Shivering Sweat Rate Sweat Loss Nylon-Spandex 26.73 C 36.87 C 38.2 W 89.3% 0.0 g/min 0.0 L Base Layer BNNT & Male 27.43 C 36.87 C 28.8 W 90.9% 0.0 g/min 0.0 L Base Layer Aerogel, BNNT 35.92 C 38.01 C 265.1 W 14.8% 14.5 g/min 0.691 L & Base Layer Nylon-Spandex 24.54 C 36.92 C 31.7 W 94.2% 0.0 g/min 0.0 L Base Layer BNNT & Female 25.84 36.91 C 24.3 W 94.2% 0.0 g/min 0.0 L Base Layer Aerogel, BNNT 35.02 C 37.94 C 176.9 W 16.5% 5.08 g/min 0.238 L & Base Layer or on a life support systems backpack. This idea has been explored in research both through NASA contract work and at the University of Colorado Boulder [7, 28, 19, 22, 75]. Chapter 2 discussed several proposals for using radiators to provide suit cooling. Still it is likely necessary to provide supplemental active cooling for extreme hot cases, especially on the Moon. A thermal protective garment like the TMG will deﬁnitely be needed during lunar EVAs to prevent large solar heat inﬂuxes that could damage the skin and body and increase the heat stress on the suit’s thermoregulation system.

An MCP suit is naturally better equipped to perform well in a Martian atmosphere. On Mars, the challenge is to provide an adequate amount of insulation to keep an astronaut warm while not building up too much evaporative resistance so that the astronaut can easily sweat during intense activity. This was shown in the simu-

114 lation. Only with proper insulation did the manikin avoid becoming very cold. But when adequate insulation was added to the suit, the manikin was able to remain at a mostly comfortable temperature throughout the EVA and also reject heat through perspiration when the activity level became strenuous. The addition of convective cooling also lowers the amount of heat that needs to be dissipated directly through sweat evaporation.

Another key takeaway is the diﬀerence in thermal results between the male and female astronaut manikins. In both the lunar and Martian environments, the female astronaut experienced lower skin temperatures, meaning increased cold discomfort. Future designs of the BioSuitTM or similar mechanical counterpressure spacesuits should include thermal and radiation protection in the list of design diﬀerences that should be made between male and female suits. Based on these results, female suits will need more thermal insulation than male suits. The sweat rates of the female astronaut were also less than the male astronaut. This means that a female suit may need to accommodate less drinking water, but this is not a certainty. The female astronaut manikins may have also sweat less than the male manikins because they were experiencing colder skin temperatures.

These results provide an interesting comparison to the results derived by Vad- havkar in his TAIThermTM Martian EVA modeling. Vadhavkar’s suit model (based on the Space Shuttle ACES ﬂight suit) had much higher thermal and evaporative resistance properties. This caused the astronaut to experience higher levels of heat stress and lower rates of evaporative cooling than in the modeling of this thesis [65]. The astronaut manikins in this thesis work experienced more cold discomfort but maintained healthier core body temperatures and were able to achieve more cooling from sweating. We believe the spacesuit model used in these simulations is a more accurate representation of a true mechanical counterpressure suit.

Overall, an MCP suit architecture like the BioSuitTM should not be limited in its implementation by thermal requirements. The capacity for cooling via both perspiration and body radiation open up a wider possibility of methods for thermal regulation. Additionally, the material properties of the clothing pieces modeled here

115 support their usage in future spacesuit testing. The Boron Nitrite Nanotube (BNNT) infused fabric provided additional radiation shielding to the chest without inhibiting moisture transport. Additionally, the Aerogel layer proved to be a necessary addition to provide insulation to an astronaut. Without this layer, the astronaut was uncomfortably cold for the majority of the EVA duration on Mars and uncomfortably hot on the Moon. As long as these insulation layers can be engineered to have a low evaporative resistance, they are an excellent candidate to keep an astronaut warm but allow for heat dissipation under high activity loads.

116 Chapter 5

Conclusions

As discussed in Chapter 1, there is a lack of knowledge as to the requirements of a life support system for a mechanical counterpressure spacesuit such as the MIT BioSuitTM. This thesis presented two primary contributions to spacesuit life support system engineering: a prototype breathing bladder for an MCP spacesuit and suit thermal modeling of planetary EVA. Both of these contributions advance the progress of MCP spacesuit development. The breathing bladder prototype work provides an example of how modern design tools, analysis and manufacturing techniques can eventually be applied to produce components of a fully functional BioSuitTM. The thermal modeling work in TAIThermTM showed that perspiration and evaporative cooling can be employed in a mechanical counterpressure suit to reduce the life support system mass by replacing the Liquid Cooling and Ventilation Garment. The modeling also demonstrated the potential of new materials to provide thermal and radiation protection in an MCP suit and highlighted the diﬀerences in thermal requirements between female and male astronauts.

117 5.1 Contributions

5.1.1 BioSuit Breathing BladderTM

The work of this thesis advances the development of the BioSuitTM project through both physical prototyping and systems level modeling. The breathing bladder prototype was built off of the designs by Webb for the Space Activity Suit and the conceptual designs of Kracik [38, 46]. This project improved upon the breathing bladders developed by Webb by introducing modern CAD capabilities as well as manufacturing flexibility enabled by the expertise at the Dainese D-Air LabTM. The result was a functional product that brought to life a part of the BioSuitTM conceptualized by Kracik. Though the unit was not a product with enough fidelity to complete human subject testing, it performed adequately in a laboratory environment, properly in- flating underneath a compressive restraining layer as part of a helmet airflow system. Overall chest contact pressure application did not meet the goal, but individual spots on the torso did reach desired pressures. The shortcomings are explained by a number of factors in the test setup including air leaks, inadequate restraint from the elastic rubber bands used for compression, and a slightly undersized mannequin torso. The physical bladder as well as it’s design and fabrication process can be used to inform development of future prototypes of BioSuitTM garment components.

5.1.2 MCP EVA Thermal Modeling

The MCP spacesuit thermal modeling yielded several important insights. Firstly that perspiration is a realistic method of heat rejection in both lunar and Martian EVA scenarios. Over 200 watts was dissipated from the spacesuit through evaporation during periods of intense activity. The remaining heat was rejected via radiation from the suit. This suggests that a mechanical counterpressure suit can achieve adequate cooling during EVA by using moisture-wicking fabrics and by incorporating several high-emissivity radiator surfaces. This has the potential to signiﬁcantly reduce the life support system mass required for temperature control. The simulations showed that

118 Boron Nitrite Nanotube (BNNT) and Aerogel infused fabrics are promising candidate materials for thermal and radiation protection. In the simulations, the astronaut manikins were frequently uncomfortably cold rather than overheated. This means that future suit engineers will need to make sure the suit has adequate insulation, especially for Mars EVAs. For lunar EVAs, radiation protection is crucial so that the astronaut does not become overheated. Additionally, the different simulation responses between male and female manikins show that thermal requirements may differ between the sexes. The female manikin had mean skin temperatures about 0.5°C colder than the male manikin and experienced more intense cold discomfort for each given suit configuration. An MCP suit like the BioSuitTM must already be tailor-fit to the wearer, but the thickness of the thermal and radiation protection layers should also depend on the individual astronaut.

5.2 Challenges and Limitations

As with any research project, there are several limitations to the conclusions that can be drawn from this thesis. The first limitation is the extensibility of the BioSuitTM breathing bladder. The manufactured bladder and helmet setup was only a prototype and is not functional for human use without significant modification. The most obvious need is for a better helmet and neck ring. The prototype bladder was molded to a wooden ring. The helmet was a simple polycarbonate plastic dome glued onto the wooden ring from the other side with a rubber gasket in between. This created an almost airtight seal, but is a permanent installation, meaning that the bladder, neck ring and helmet are all attached as one piece. A design with a metal neck ring and detachable locking helmet is needed for human subject testing. The test setup had several limitations as well. The air compressor pump could not sustain the desired airflow rate of 6 cubic feet per minute (0.17 liters per minute). Several leaks existed in the system as well. The rubber bands used to provide counterpressure provided inconsistent pressure and a better method of compression is needed. The bladder was also oversized for the mannequin used and it was difficult to adhere

119 the pressure sensors directly onto the chest surface. Both of these factors mean that the pressure sensor data readings were inconsistent and the data collected was only considered pilot data. The thermal model built in TAIThermTM also has several limitations. The most prominent of these are the material properties of the suit layers applied to the manikin. Key variables like whole body thermal resistance could not be found for the fabric layers since all suit layers are still the subject of current research. Values such as whole body thermal resistance and and whole body evaporative resistance are often determined on an individual garment basis using laboratory testing. No such testing was done before this modeling. Values were assigned based on analogue clothing materials and garments, but there are likely to be discrepancies between the real material properties and those used in the TAITherm simulations. The weather files may also be slightly inaccurate. We were unable to find values for several input variables and therefore made decisions based off of Earth analogues and educated guesses.

5.3 Future Work

There remains much unﬁnished and unexplored work related to this research. The BioSuitTM as a whole remains a concept, even after almost two decades of investigation. Design, materials selection, fabrication methods, and prototype development and testing of garment and life support components remain areas of ongoing work. For the BioSuitTM breathing bladder, future work includes testing the contact pressure changes associated with a chest volume change. Maintaining a constant applied pressure throughout the breathing cycle is of course is the true purpose of the breathing bladder. Testing could be conducted by applying an external force to the chest, or perhaps the use of a CPR mannequin with a a ﬂexible chest volume. Eventually, a complete design of a helmet system that integrates seamlessly with the external counter-lung and life support backpack will be needed. It is also an unanswered question how this bladder would interface with the outer pressure layer

120 of an MCP suit. Additionally, alternative computer modeling programs such as CAD-VIDYA that are built for soft goods rather than solid bodies should be used for designing and engineering analysis. An analysis of geometric design changes needed for female astronauts and other body types should also be considered. One area that could be studied and tested in more detail is the spatial variation in volume expansion of the chest. During quiet breathing, most volume change comes from below the sternum, while the opposite is true for heavy breathing. The design in this thesis did not consider this, but future bladders and suit prototypes may accommodate more volume change in the upper chest. As discussed, breathing bladders in past flight suits and the Space Activity Suit often caused overheating on the chest. It is possible that supplemental cooling will be needed to prevent uncomfortable perspiration build-up underneath the bladder during EVA operation. This problem needs further investigation. The long term goal for this project is testing of a breathing bladderin a closed loop air-recirculating system and eventually with a full human subject to gather objective performance data as well as subjective human feedback on garment fit and comfort. Future work can build on the modeling done by verifying and validating the material properties of the clothing layers modeled. The BNNT and Aerogel fabrics are still in early development and therefore their properties modeled in this thesis were based on the best current knowledge. Future work may reveal new information that changes the validity of the modeling results. Simulations like the ones presented in this thesis could be used to determine the exact material thicknesses or material properties needed in an actual BioSuitTM to keep an astronaut at a completely comfortable temperature throughout the EVA. A wider range of lunar and Martian environmental conditions should also be tested. The simulations could also be run for a full eight hour EVA to gauge effects that may not occur in the shorter four hour EVA timeline used in this modeling. The differences between female and male thermal responses can also be investigated further than in this thesis. This thermal model only switched the sex of the manikin

121 used in the model, but a more researched and detailed analysis should be conducted that does not rely entirely on TAITherm’s built-in functions. Even with the simple testing in this thesis it was found that female astronauts will require slightly diﬀerent conditions to remain in a comfortable thermal state during EVA while wearing MCP suits. Therefore more research is needed to determine what suit designs will make both female and male astronauts safe and comfortable during EVA. Finally, it is worth investigating whether the ability to sweat directly into the environment will violate any planetary protection policies on the Moon or Mars.

5.4 Final Conclusions

It has now been almost 50 years since Apollo 17 commander Eugene Cernan stepped back onto the Apollo Lunar Module in December of 1972 and brought the ﬁrst phase of human lunar exploration to an end. No other planetary EVAs have occurred since. NASA has announced its intention to end that draught by returning to the Moon in the late 2020s and venturing on to Mars in the 2030s [2]. To enable a sustainable expansion of human space exploration improvements must be made in a litany of technologies, but life support systems are one of the most critical. The next several decades oﬀer a very exciting opportunity to support human missions to the Moon and Mars. A mechanical counterpressure spacesuit like the BioSuitTM can aid in this mission by increasing astronaut productivity during EVAs while likely saving mass. We found that two key parts of this suit—a breathing bladder that will allow an astronaut to breathe easily and a specially designed garment that enables natural thermoregulation—are both possible to realize using modern engineering tools. The work of this thesis helps bring the next generation of planetary spacesuits one step closer to reality.

122 Appendix A

TAIThermTM EVA Human Visualizations

123 A.1 Moon EVA Simulation Results

Figure A-1: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Lunar EVA wearing only a skinsuit. The extreme solar loading is evidence by the ”hot” side of the astronaut experiencing skin temperatures of up to 60°C while the ”cold” side temperatures drop to 12°C. These eﬀects would be more equalized in an EVA in which the astronaut does not maintain a constant orientation relative to the Sun. Clearly a radiation shielding garment is needed for direct solar exposure.

124 Figure A-2: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Lunar EVA wearing a skinsuit and BNNT jacket. The jacket helps protect the torso from some of the most severe solar loading eﬀects.

125 Figure A-3: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Lunar EVA wearing the full clothing layup. Skin surface temperatures are much more uniform (lower on the hot side and higher on the cold side). However the astronaut still feels slightly chilled throughout the EVA.

126 A.2 Mars EVA Simulation Results

Figure A-4: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Martian EVA wearing only a skinsuit. The astronaut experiences increasingly cold discomfort, oﬀset only a little by the metabolically strenuous phases of the EVA.

127 Figure A-5: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Martian EVA wearing skinsuit and BNNT torso over layer. The astronaut experiences increasingly cold discomfort, oﬀset only a little by the metabolically strenuous phases of the EVA.

128 Figure A-6: Skin surface temperature (top) and Berkeley Comfort thermal sensation (bottom) for a Martian EVA wearing a skinsuit, BNNT jacket and Aerogel insulation layer. Here the astronaut maintains comfortable skin temperatures for almost the entirety of the EVA, making this suit layup the best performing of all those tested.

129 130 Appendix B

Eﬀect of Layering a Garment in TAIThermTM

131 To verify that modeling the base layer of an MCP suit as a single layer was appropriate, a test was conducted to assess whether the number of layers a garment was split into effected the thermal results. All garment configurations of the nylon- spandex base layer had the same total thickness of 4.2 mm, but this thickness was split into one, two or three layers. Modeled in FigureB-1 is a male astronaut in the Martian EVA conditions wearing the nylon-spandex base layer in different layering configurations. The introduction of layering had no significant effect on the body’s core temperature and only a 0.5°C change in mean skin temperature (MST) at the end of a 4 hour EVA. Therefore the decision was made that it was safe to model the base layer as a single layer in TAITherm

Figure B-1: Comparison of body temperatures during Martian EVA resulting from a base layer of 4.2 thickness kept as a single layer or split into two or three layers.

132 Appendix C

Digital Archive

133 C.1 Contents of Digital Archive

A digital archive of the research materials of this thesis is available on Dropbox. This folder contains original images used, testing videos, novel pliance® pressure data, MATLAB scripts for plotting data, and TAIThermTM input and output ﬁles. Read- ers may view the folder at this link: https://www.dropbox.com/sh/7jv0150adeocrcq/AADNWCWb8Us885xGdBUXfo5Ea?dl=0.

To gain access please contact the author at [email protected] or Professor Dava Newman at [email protected].

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