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After several decades of human spaceflight, the community of space-faring nations has accumulated a diverse and sometimes harrowing history of toxicological events that have plagued human space endeavors almost from the very beginning. Some lessons have been learned in ground-based test beds and others were discovered the hard way - when human lives were at stake in space. From such lessons one can build a risk-management framework for toxicological events to minimize the probability of a harmful exposure, while recognizing that we cannot predict all possible events. Space toxicologists have learned that relatively harmless compounds can be converted by air revitalization systems into compounds that cause serious harm to the crew.

The objective of this study is to evaluate ventilation efficiency regarding to the International Space Station (ISS) cabin ventilation during the ISS assembly mission 1J. The focus is on carbon dioxide spatial/temporal variations within the Node 2 and attached modules. An integrated model for CO2 transport analysis that combines 3D CFD modeling with the lumped parameter approach has been implemented. CO2 scrubbing from the air by means of two ISS removal systems is taken into account. It has been established that the ventilation scheme with an ISS Node 2 bypass duct reduces short-circuiting effects and provides less CO2 gradients when the Space Shuttle Orbiter is docked to the ISS. This configuration results in reduced CO2 level within the ISS cabin.

Boeing has contracts for military application of twin engine airplanes generically identified in this paper as the MX airplane. Unlike previous derivatives, the MX airplanes are produced with a streamlined manufacturing process to improve cost and schedule performance. The final assembly of each MX airplane includes a series of integration tests, called factory functional tests (FFTs), which are modified from those of typical commercial versions and verify correctness of equipment installation and basic functionalities. Two airplanes have been through the production line resulting in a number of FFT lessons learned. Addressed are the power distribution lessons learned: 1) the expanded coverage of the basic automated power-on generation system test, 2) the need for a manual wire continuity test, 3) salient features of the power distribution tests, and 4) keys to make first pass power distribution test smooth and successful.

NASA's proposed lunar lander, Altair, will be exposed to vastly different external temperatures following launch till its final destination on the moon. In addition, the heat rejection is lowest at the lowest environmental temperatures (0.5 kW @ 4K) and highest at the highest environmental temperature (4.5 kW @ 215K). This places a severe demand on the radiator design to handle these extreme turn-down requirements. A radiator with digital turn-down capability is currently under study at JPL as a robust means to meet the heat rejection demands and provide freeze protection while minimizing mass and power consumption. Turndown is achieved by independent control of flow branches with isolating latch valves and a gear pump to evacuate the isolated branches. A bench-top test was conducted to characterize the digital radiator concept. Testing focused on the demonstration of proper valve sequencing to achieve turn-down and recharge of flow legs.

Environmental Control Systems (ECS) ducts on airplanes are primarily fabricated from aluminum or thermoset composites, depending on temperature and pressure requirements. It is imperative to fabricate lightweight, cost effective, durable, and repairable systems with minimal tooling. It is also important that the duct systems are easy to assemble even with alignment issues resulting from structural variations, tolerance accumulation, variation from thermal expansion of different materials, and inherent duct stiffness. These requirements create an opportunity and need for a technology that can address all of these issues, while increasing performance at the same time. This report provides a background on current ECS ducting systems.

Solar proton events (SPEs) represent the single-most significant source of acute radiation exposure during space missions. Historically, an exponential in rigidity (particle momentum) fit has been used to express the SPE energy spectrum using GOES data up to 100 MeV. More recently, researchers have found that a Weibull fit better represents the energy spectrum up to 1000 MeV (1 GeV). In addition, the availability of SPE data extending up to several GeV has been incorporated in analyses to obtain a more complete and accurate energy spectrum representation. In this paper we discuss the major SPEs that have occurred over the past five solar cycles (~50+ years) in detail - in particular, Aug 1972 and Sept & Oct 1989 SPEs. Using a high-energy particle transport/dose code, radiation exposure estimates are presented for various thicknesses of aluminum. The effects on humans and spacecraft systems are also discussed in detail.

The System Architecture Virtual Integration (SAVI) program is a collaboration of industry, government, and academic organizations within the Aerospace Vehicle System Institute (AVSI) with the goal of structuring a new integration process that relies on a “single-truth” architectural framework. The SAVI approach of “Integrate, then Build” provides a modern distributed development environment which arrests the propagation of requirements errors through the development life cycle. It does so by capturing design assumptions and shared properties of the system design in an authoritative, annotated architectural model. This reference model provides a common, analyzable framework for confirming that system requirements remain complete, consistent, and correct at all levels of system decomposition. Core concepts of SAVI include extensive use of model-based system engineering tools and use of a “single-truth” reference architectural model.

The power and thermal management system (PTMS) developed by Honeywell for aircraft is an integral approach combining the functions of the auxiliary power unit (APU), emergency power unit (EPU), environmental control system (ECS), and thermal management system (TMS). The next generation PTMS discussed in this paper incorporates the new more electric architecture (MEA) and energy efficient aircraft (EEA) initiatives. Advanced system architectures with increased functionality and further integration capabilities with other systems are included. Special emphasis is given to improvements resulting from interactions with the main engine, main electric power generation, and flight actuation. The major drivers for advancement are highlighted, as well as the potential use of new technologies for turbomachinery, heat exchangers, power electronics, and electric machines. More advanced control and protection algorithms are considered.

Launch vehicles that use electric actuators for thrust vector or flight control require a safe, reliable and lightweight source of electrical power. Honeywell, working with NASA Glenn Research Center and Lockheed Martin Space Systems, has developed and successfully tested a turbine-driven electric power generation system which meets these needs. This Turbine Power Unit (TPU) uses hydrogen and oxygen propellants which react catalytically to drive a shaft-speed turboalternator mounted on foil bearings. A high-reactance permanent-magnet machine (HRPMM) was selected for this application. The power conditioning and control electronics can be located within the TPU housing and the hydrogen fuel can be used to pressurize the bearings and electronics and to regeneratively cool the machine. A brassboard unit incorporating many of these features was successfully tested at output power levels from 0 to 138 kilowatts (kW).

Spacecraft radiators reject heat to their surroundings. Radiators can be deployable or mounted on the body of the spacecraft. NASA's Crew Exploration Vehicle is to use body mounted radiators. Coatings play an important role in heat rejection. The coatings provide the radiator surface with the desired optical properties of low solar absorptance and high infrared emittance. These specialized surfaces are applied to the radiator panel in a number of ways, including conventional spraying, plasma spraying, or as an appliqué. Not specifically designed for a weathering environment, little is known about the durability of conventional paints, coatings, and appliqués upon exposure to weathering and subsequent exposure to solar wind and ultraviolet radiation exposure. In addition to maintaining their desired optical properties, the coatings must also continue to adhere to the underlying radiator panel.

For NASA's Constellation Program, an Intravehicular Activity (IVA) and Extravehicular Activity (EVA) Life Support Umbilical System (LSUS) will be required to provide environmental protection to the suited crew during Crew Exploration Vehicle (CEV) cabin contamination or depressurization and contingency EVAs. The LSUS will provide the crewmember with ventilation, cooling, power, communication, and data, and will also serve as a crew safety restraint during contingency EVAs. The LSUS will interface with the Vehicle Interface Assembly (VIA) in the CEV and the Suit Connector on the suit. This paper describes the effort performed to develop concept designs for IVA and EVA umbilicals, universal multiple connectors, handling aids and stowage systems, and VIAs that meet NASA's mission needs while adhering to the important guiding principles of simplicity, reliability, and operability.

Carbon dioxide (CO2) removal technology development for portable life support systems (PLSS) has traditionally concentrated in the areas of solid and liquid chemical sorbents and semi-permeable membranes. Most of these systems are too heavy in gravity environments, require prohibitive amounts of consumables for operation on long term planetary missions, or are inoperable on the surface of Mars due to the presence of a CO2 atmosphere. This paper describes the effort performed to mature an innovative CO2 removal technology that meets NASA's planetary mission needs while adhering to the important guiding principles of simplicity, reliability, and operability. A breadboard cryogenic carbon dioxide scrubber for an ejector-based cryogenic PLSS was developed, designed, and tested. The scrubber freezes CO2 and other trace contaminants out of expired ventilation loop gas using cooling available from a liquid oxygen (LOX) based PLSS.

Providing water necessary to maintain life support has been accomplished in spacecraft vehicles for over forty years. This paper will investigate how previous U.S. space vehicles provided potable water. The water source for the spacecraft, biocide used to preserve the water on-orbit, water stowage methodology, materials, pumping mechanisms, on-orbit water requirements, and water temperature requirements will be discussed. Where available, the hardware used to provide the water and the general function of that hardware will also be detailed. The Crew Exploration Vehicle (CEV or Orion) water systems will be generically discussed to provide a glimpse of how similar they are to water systems in previous vehicles. Conclusions, questions, and recommendations on strategies that could be applied to CEV based on previous spacecraft water system lessons learned will be made.

As part of preparing for the Crew Exploration Vehicle (CEV), the National Aeronautics and Space Administration (NASA) worked on developing the requirements to manage the fire risk. The new CEV poses unique challenges to current fire protection systems. The size and configuration of the vehicle resembles the Apollo capsule instead of the current Space Shuttle or the International Space Station. The smaller free air volume and fully cold plated avionic bays of the CEV requires a different approach in fire protection than the ones currently utilized. The fire protection approach discussed in this paper incorporates historical lessons learned and fire detection and suppression system design philosophy spanning from Apollo to the International Space Station.

The Carbon Dioxide Removal Assembly (CDRA) is a part of the International Space Station (ISS) Environmental Control and Life Support (ECLS) system. The CDRA provides carbon dioxide (CO2) removal from the ISS on-orbit modules. Currently, the CDRA is the secondary removal system on the ISS, with the primary system being the Russian Vozdukh. Within the CDRA are two Desiccant/Adsorbent Beds (DAB), which perform the carbon dioxide removal function. The DAB adsorbent containment approach required improvements with respect to adsorbent containment. These improvements were implemented through a redesign program and have been implemented on units on the ground and returning from orbit. This paper presents a DAB design modification implementation description, a hardware performance comparison between the unmodified and modified DAB configurations, and a description of the modified DAB hardware implementation into the on-orbit CDRA.

Both humans and onboard radiosensitive systems (electronics, materials, payloads and experiments) are exposed to the deleterious effects of the harsh space radiations found in the space environment. The purpose of this paper is to present the space radiation environment extended to deep space based on environment models for the moon, Mars, Jupiter, and Saturn and compare these radiation environments with the earth's radiation environment, which is used as a comparative baseline. The space radiation environment consists of high-energy protons and electrons that are magnetically “trapped” in planetary bodies that have an intrinsic magnetic field; this is the case for earth, Jupiter, and Saturn (the moon and Mars do not have a magnetic field). For the earth this region is called the “Van Allen belts,” and models of both the trapped protons (AP-8 model) and electrons (AE-8 model) have been developed.

Similar to finding a home on Earth, location is important when selecting where to set up an exploration outpost. Essential considerations for comparing potential lunar outpost locations include: (1) areas nearby that would be useful for In-Situ Resource Utilization (ISRU) oxygen extraction from regolith for crew breathing oxygen as well as other potential uses; (2) proximity to a suitable landing site; (3) availability of sunlight; (4) capability for line-of-sight communications with Earth; (5) proximity to permanently-shadowed areas for potential in-situ water ice; and (6) scientific interest. The Mons Malapert1 (Malapert Mountain) area (85.5°S, 0°E) has been compared to these criteria, and appears to be a suitable location for a lunar outpost.

Currently scheduled to be delivered to the International Space Station (ISS) in 2009, Crew Quarters (CQs) will be installed in the Node 2 Module. The CQs provide crewmembers with private space, a place to sleep, and minimal storage. Analysis is to be performed to determine if the United States Operational Segment (USOS) Node 2 can maintain temperature between 47°C and 62°C (65°F and 80°F) [units are CCGS with U.S unit in parenthesis] within the CQ. The analysis will concentrate on the nominal hot environmental case. Environmental heat is due to solar heating of the external shell of the ISS. Configurations including both three and four CQs are examined, as well as multiple configurations of the Low Temperature Loop (LTL) that flows through the Node 2 Common Cabin Air Assembly (CCAA). This paper describes the analysis performed to determine if Node 2 will be able to maintain cabin temperature between 47°C and 62°C (65°F and 85°F).

A procedure for calculating the engine inlet diffuser section liquid water content and mass fractions of liquid water content associated with the water droplet size distribution for wind milling engine ice accretion analysis is presented. Critical fuel reserve calculation for extended twin-engine operation requires the determination of drag increase due to ice accretion on inoperative wind milling engine fan blade and guide vane.

There is an increasing demand in the aerospace industry for automated machinery that is portable, flexible and light. This paper will focus on a joint project between BROETJE-Automation and Boeing called the Universal Splice Machine (USM). The USM is a portable, flexible and lightweight automated drilling and fastening machine for longitudinal splices. The USM is the first machine of its kind that has the ability not only to drill holes without the need to deburr, (burrless drilling) but also to insert fasteners. The Multi Function End Effector (MFEE) runs on a rail system that is mounted directly on the fuselage using a vacuum cup system. Clamp up is achieved through the use of an advanced electromagnet. A control cart follows along next to the fuselage and includes an Automated Fastener Feeding System. This paper will show how this new advancement has the capabilities to fill gaps in aircraft production that automation has never reached before.