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In order to provide effective cooling for astronauts during extravehicular activities (EVAs), a liquid cooling and ventilation garment (LCVG) is used to remove heat by a series of tubes through which cooling water is circulated. To better predict the effectiveness of the LCVG and determine possible modifications to improve performance, computer simulations dealing with the interaction of the cooling garment with the human body have been run using the Wissler Human Thermal Model. Simulations have been conducted to predict the heat removal rate for various liquid cooled garment configurations. The current LCVG uses 48 cooling tubes woven into a fabric with cooling water flowing through the tubes. The purpose of the current project is to decrease the overall weight of the LCVG system. In order to achieve this weight reduction, advances in the garment heat removal rates need to be obtained.

For planetary exploration, new thermal insulation materials are needed to deal with unique environmental conditions presented to extravehicular activity (EVA). The thermal insulation material and system used in the existing space suit were specifically designed for low orbit environment. They are not adequate for low vacuum condition commonly found in planetary environments with a gas atmosphere. This study attempts to identify the types of lofty nonwoven thermal insulation materials and the construction parameters that yield the best performance for such application. Lofty nonwovens with different construction parameters are evaluated for their thermal conductivity performance. Three different types of fiber material: solid round fiber, hollow fiber, and grooved fiber, with various denier, needling intensity, and web density were evaluated.

A thermal analysis of the compressible carbon dioxide (CO2) flow for the Portable Fire Extinguisher (PFE) system has been performed. A SINDA/FLUINT model has been developed for this analysis. The model includes the PFE tank and the Temporary Sleep Station (TeSS) nozzle, and both have an initial temperature of 72 °F. In order to investigate the thermal effect on the nozzle due to discharging CO2, the PFE TeSS nozzle pipe has been divided into three segments. This model also includes heat transfer predictions for PFE tank inner and outer wall surfaces. The simulation results show that the CO2 discharge rates and component wall temperatures fall within the requirements for the PFE system. The simulation results also indicate that after 50 seconds, the remaining CO2 in the tank may be near the triple point (gas, liquid and solid) state and, therefore, restricts the flow.

The NASA-wide CEV Smart Buyer Team (SBT) was assembled in January 2006 and was tasked with the development of a NASA in-house design for the CEV Crew Module (CM), Service Module (SM), and Launch Abort System (LAS). This effort drew upon over 250 engineers from all of the 10 NASA Centers. In 6 weeks, this in-house design was developed. The Thermal Systems Team was responsible for the definition of the active and passive design architecture. The SBT effort for Thermal Systems can be best characterized as a design architecting activity. Proof-of-concepts were assessed through system-level trade studies and analyses using simplified modeling. This nimble design approach permitted definition of a point design and assessing its design robustness in a timely fashion. This paper will describe the architecting process and present trade studies and proposed thermal designs

Thermal control systems for space application plant growth chambers offer unique challenges. The ability to control temperature and humidity independently gives greater flexibility for optimizing plant growth. Desired temperature and relative humidity range vary widely from 15°C to 35°C and 65% to 85% respectively. On top of all of these variables, the thermal control system must also be conservative in power and mass. These requirements to develop and test a robust thermal control system for space applications led to the design and development of the Environmental System Test Stand (ESTS) at NASA Johnson Space Center (JSC). The ESTS was designed to be a size constrained, environmental control system test stand with the flexibility to allow for a variety of thermal and lighting technologies. To give greater understanding to the environmental control system, the development of the ESTS included both mathematical models and the physical test stand.

Research committed by the Langley Research Center through 1995 resulting in the HZETRN code provides the current basis for shield design methods according to NASA STD-3000 (2005). With this new prominence, the database, basic numerical procedures, and algorithms are being re-examined with new methods of verification and validation being implemented to capture a well defined algorithm for engineering design processes to be used in this early development phase of the Bush initiative. This process provides the methodology to transform the 1995 HZETRN research code into the 2005 HZETRN engineering code to be available for these early design processes. In this paper, we will review the basic derivations including new corrections to the codes to insure improved numerical stability and provide benchmarks for code verification.

Meeting radiation protection requirements during EVA is predominantly an operational issue with some potential considerations for temporary shelter. The issue of spacesuit shielding is mainly guided by the potential of accidental exposure when operational and temporary shelter considerations fail to maintain exposures within operational limits. In this case, very high exposure levels are possible which could result in observable health effects and even be life threatening. Under these assumptions, potential spacesuit radiation exposures have been studied using known historical solar particle events to gain insight on the usefulness of modification of spacesuit design in which the control of skin exposure is a critical design issue and reduction of blood forming organ exposure is desirable.

Major improvements have recently been completed in the approach to spacecraft shielding analysis. A Computer-Aided Design (CAD)-based system has been developed for determining the shielding provided to any point within or external to the spacecraft. Shielding analysis is performed using a commercially available stand-alone CAD system and a customized ray-tracing subroutine contained within a standard engineering modeling software package. This improved shielding analysis technique has been used in several vehicle design projects such as a Mars transfer habitat, pressurized lunar rover, and the redesigned Space Station. Results of these analyses are provided to demonstrate the applicability and versatility of the system.

Automatic thermal comfort control for the minimum consumables PLSS is undertaken using several control approaches. Accuracy and performance of the strategies using feedforward, feedback, and gain scheduling are evaluated through simulation, highlighting their advantages and limitations. Implementation issues, consumable usage, and the provision for the extension of these control strategies to the cryogenic PLSS are addressed.

The Self Contained Atmospheric Protective Ensemble (SCAPE) suits, worn at the Kennedy Space Center (KSC) have been updated from the original 1970's design. The suits were renamed Propellant Handlers Ensemble (PHE) but are still commonly referred to as SCAPE. Several modifications to the suit were done over the last 20 years to improve the design for operational use. However, anthropometric changes in the user population over time have not been addressed. The following study addressed anthropometric concerns in the current SCAPE population. It was found that all suits had at least one area in which the recommended upper limit was exceeded by technicians. The most common areas to exceed the upper limit were: waist circumference, chest circumference and upper thigh circumference. Forearm circumference posed the least concern unless using long gauntlet glove which cause the twist lock ring to be located at the forearm rather than the wrist.

The NASA Johnson Space Center has plans to integrate a Solid Amine Water Desorbed (SAWD II) carbon dioxide removal subsystem into the Variable Pressure Growth Chamber (VPGC). The SAWD II subsystem will be used to remove any excess carbon dioxide (CO2) input into the VPGC which is not assimilated by the plants growing in the chamber. An analysis of the integrated VPGC-SAWD II system was performed using a mathematical model of the system implemented in the Computer-Aided System Engineering and Analysis (CASE/A) package. The analysis consisted of an evaluation of the SAWD II subsystem configuration within the VPGC, the planned operations for the subsystem, and the overall performance of the subsystem and other VPGC subsystems. Based on the model runs, recommendations were made concerning the SAWD II subsystem configuration and operations, and the chambers' automatic CO2 injection control subsystem.

Biological wastewater processing has been under investigation by AlliedSignal Aerospace and NASA Johnson Space Center (JSC) for future use in space. Testing at JSC in the Hybrid Regenerative Water Recovery System (HRWRS) in preparation for future closed human testing has been performed. Computer models have been developed to aid in the design of a new four-person immobilized cell bioreactor. The design of the reactor and validation of the computer model is presented. In addition, the total organic carbon (TOC) computer model has been expanded to begin investigation of nitrification. This model is being developed to identify the key parameters of the nitrification process, and to improve the design and operating conditions of nitrifying bioreactors. In addition, the model can be used as a design tool to rapidly predict the effects of changes in operational conditions and reactor design, significantly reducing the number and duration of experiments required.

This paper present a summary of the design studies for the suit port proof of concept. The Suit Port reduces the need for airlocks by docking the suits directly to a rover or habitat bulkhead. The benefits include reductions in cycle time and consumables traditionally used when transferring from a pressurized compartment to EVA and mitigation of planetary surface dust from entering into the cabin. The design focused on the development of an operational proof of concept evaluated against technical feasibility, level of confidence in design, robustness to environment and failure, and the manufacturability. A future paper will discuss the overall proof of concept and provide results from evaluation testing including gas leakage rates upon completion of the testing program.

Vibrational disturbance magnitude and frequency on space-flight missions is often a critical factor regarding mission success. Both materials processing experiments and astronomical investigations have specific microgravity environmental requirements. Recent efforts have been made to quantify the microgravity environment on the Space Shuttle Columbia by measuring gravity levels produced by specific mission events such as Orbiter engine burns, treadmill and ergometer activities, crew sleep periods, rotating chair operations, and body mass measurement operations. However, no measurements have been made of specific, nominal crewmember activities such as translating about the middeck, flight-deck, or in the Spacelab. This report presents pilot-study data of test subject forces induced by intravehicular activities such as push-offs and landings with both hands and feet. Five subjects participated in this investigation.

An ADvanced Automotive Manikin (ADAM) developed at the National Renewable Energy Laboratory (NREL) is used to evaluate NASA’s liquid cooling garments (LCGs) used in advanced spacesuits. The manikin has 120 separate heated/sweating zones and is controlled by a finite-element physiological model of the human thermo-regulatory system. Previous testing showed the thermal sensation and comfort followed expected trends as the LCG inlet fluid temperature was changed. The Phase II test data demonstrates the repeatability of ADAM by retesting the baseline LCG. Skin and core temperature predictions using ADAM in an LCG/arctic suit combination are compared to NASA physiological data to validate the manikin/model. An additional Orlan LCG configuration is assessed using the manikin and compared to the baseline LCG.

The incidence of DCS in null gravity appears to be considerably less than predicted by 1-g experiments. In NASA studies in 1-g, 83% of the incidents of DCS occur in the legs. We report first on a study with a crossover design that indicated a considerable reduction in the decompression Doppler bubble grade in the lower extremities in subjects in simulated microgravity (bed rest) as compared to themselves when ambulatory in unit gravity. Second we describe the results of a cardiovascular deconditioning study using a tail-suspended rat model. Since there may be a reduction in bubble production in 0-g, this would reduce the possibility of acquiring neurological DCS, especially by arterial gas embolism. Further, cardiovascular deconditioning appears to reduce the pulmonary artery hypertension (secondary to gas embolization) necessary to effect arterialization of bubbles.

Gas-separation and reverse-osmosis membrane models are being developed in conjunction with membrane testing at NASA JSC. The completed gas-separation membrane model extracts effective component permeabilities from multicomponent test data, and predicts the effects of flow configuration, operating conditions, and membrane dimensions on module performance. Variable feed- and permeate-side pressures are considered. The model has been applied to test data for hollow-fiber membrane modules with simulated cabin-air feeds. Results are presented for a membrane designed for air drying applications. Extracted permeabilities are used to predict the effect of operating conditions on water enrichment in the permeate. A first-order reverse-osmosis model has been applied to test data for spiral wound membrane modules with a simulated hygiene water feed. The model estimates an effective local component rejection coefficient under pseudo-steady-state conditions.

This paper describes the need and solution for minimizing EVA airlock time and depress gas losses using a new method that minimizes EVA out-the-door time for a suited astronaut and reclaims most of the airlock depress gas. This method consists of one or more related concepts that use an evacuated reservoir tank to store and reclaim the airlock depress gas. The evacuated tank can be an inflatable tank, a spent fuel tank from a lunar lander descent stage, or a backup airlock. During EVA airlock operations, the airlock and reservoir would be equalized at some low pressure, and through proper selection of reservoir size, most of the depress gas would be stored in the reservoir for later reclamation. The benefit of this method is directly applicable to long duration lunar and Mars missions that require multiple EVA missions (up to 100, two-person lunar EVAs) and conservation of consumables, including depress pump power and depress gas.

One of the risks associated with long-term space flights is cancer incidence resulting from chronic exposure to space radiation. Assessment of incurred risk from radiation exposure requires quantifying the dose throughout the body. The space radiation exposure received by Space Shuttle astronauts is measured by thermoluminescent dosimeters (TLDs) worn during every mission. These dosimeters measure the absorbed dose to the skin, but the dose to internal organs is required for estimating the cancer risk induced by space radiation. A method to extrapolate these skin dose measurements to realistic organ specific dose estimates, using the Computerized Anatomical Man (CAM) and Computerized Anatomical Female (CAF) models, is discussed in detail. A transport code, which propagates high energy nucleon and charged particles, is combined with the CAM/CAF-generated shielding areal distributions to evaluate the absorbed dose at selected organ sites.

The quantitative analysis of the food system for long-term space missions is a crucial factor for the comparison of different food plans and for the evaluation of the food system as part of the overall mission. Such analysis should include important factors such as nutrition, palatability, diet cycle length, and psychological issues related to food. This paper will give the details of a mathematical model that was developed during the first author's participation as a Summer Faculty Fellow at Johnson Space Center. The model includes nutrition, palatability, diet cycle length, and psychological issues as important components. The model is compatible with the Equivalent System Mass (ESM) metric previously developed as the Advance Life Support (ALS) Research and Technology Metric.