Search Results

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.

Development of a second generation Electronic Nose at JPL is focusing on optimization of the sensing films to increase sensitivity and optimization of the array. Toward this goal, studies have focused on sources of noise in the films, alternatives to carbon black as conductive medium, measurement techniques, and development of an analytical approach to polymer selection to maximize the abilities of the array to distinguish among compounds.

This is Part 3 of a development program to evaluate candidate nonablative aeroshell designs. The primary goal of this C/C aeroshell development task was to demonstrate the feasibility and performance of a lightweight C/C non-ablative aeroshell design that integrates advanced C/C materials and structural configurations. The thermal performance was evaluated by Arc Jet testing at NASA Ames of representative structural models. In this phase of the program, new carbon-carbon materials and structural core designs were evaluated, as well as an alternative aerogel material. The test models were composed of a quasi-isotropic Carbon/Carbon(C/C) front face sheet (F/S), eggcrate or honeycomb core, C/C back F/S, Carbon and resorcinol-formaldehyde aerogel insulation. Part One of this work [1] demonstrated the feasibility through arc-jet testing and Part Two [2] included analytical modeling of the test geometry to validate the design.

Candidate Aeroshell Test models composed of a quasi-isotropic Carbon/Carbon(C/C) front face sheet (F/S), eggcrate core, C/C back F/S, Carbon Aerogel insulation, C/C radiation shield and the C/C close-out were constructed based on the analytical temperature predictions presented in Part One of this work[1]. The analytical results obtained for a simulated Mars entry of a 2.9 meter diameter cone shaped Carbon-Carbon Aeroshell demonstrated the feasibility of the design. These results showed that the maximum temperature the front F/S reached during the decent was 1752 °C with the resulting rear temperature reaching 326 °C in the thermal model. Part Two of this work documents the thermal modeling and correlation for the Mars Aeroshell test sample and fixture. A finite difference, SINDA/G, thermal math model of the test fixture and sample was generated and correlated to data from an arc jet test conducted at the NASA Ames Research Center's interactive heating facility.

The Tropospheric Emission Spectrometer (TES) is a cryogenic instrument which will be launched on NASA's Earth Observation System (EOS) Chemistry Platform in the year 2003. The overall mission lifetime for the instrument is 5 years with an additional period of 2 years required for ground test and calibration. The EOS Chemistry Platform will be placed in a sun-synchronous near-circular polar orbit with an inclination of 98.2 degrees and a mean altitude of 705 km. The overall objective of TES is the investigation and quantification of global climate change, both natural and anthropogenic. It is a high resolution infrared imaging (1×16 pixels) Fourier Transform Spectrometer with spectral coverage of 3.3-15.4 μm at a spectral resolution of 0.10 cm−1 or 0.025 cm−1 intended for the measurement and profiling of essentially all infrared-active molecules present in the Earth's lower atmosphere (0-30+ km).

The Multi-angle Imaging SpectroRadiometer (MISR) instrument was launched aboard NASA’s Earth Observing System (EOS) Terra spacecraft on December 18, 1999. The overall mission design lifetime for the instrument is 6 years. The EOS Terra spacecraft was placed in a sun-synchronous near-circular polar orbit with an inclination of 98.3 degrees and a mean altitude of 705 km. The overall objective of MISR is to provide a means to study the ecology and climate of Earth through the acquisition of global multiangle imagery on the daylit side of Earth. MISR views the sunlit Earth simultaneously at nine widely spaced angles, collects global images with high spatial detail in four colors at every angle. The images acquired, once calibrated, provide accurate measurements of brightness, contrast and color of reflected sunlight.

This paper presents the analysis findings of a study reducing the overall mass of the lightweight liquid cooling garment (LCG). The LCG is a garment worn by crew to actively cool the body, for spacesuits and launch/entry suits. A mass reduction of 66% was desired for advanced missions. A thermal math model of the LCG was developed to predict its performance when various mass-reducing changes were implemented. Changes included varying the thermal conductivity and thickness of the garment or of the coolant tubes servicing the garment. A second model was developed to predict behavior of the suit when the cooling tubes were to be removed, and replaced with a highly-conducting (waterless) material. Findings are presented that show significant reductions in weight are theoretically possible by improving conductivity in the garment material.

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

Hamilton Sundstrand has developed a scalable evaporative heat rejection system called the Multi-Fluid Evaporator (MFE). It was designed to support the Orion Crew Module and to support future Constellation missions. The MFE would be used from Earth sea level conditions to the vacuum of space. This system combines the functions of the Space Shuttle flash evaporator and ammonia boiler into a single compact package with improved freeze-up protection. The heat exchanger core is designed so that radial flow of the evaporant provides increasing surface area to keep the back pressure low. The multiple layer construction of the core allows for efficient scale up to the desired heat rejection rate. A full-scale unit uses multiple core sections that, combined with a novel control scheme, manage the risk of freezing the heat exchanger cores. A four-core MFE prototype was built in 2007.

Hamilton Sundstrand is under contract with the NASA Johnson Space Center to develop a scalable, evaporative heat rejection system called the Multi-Fluid Evaporator (MFE). It is being designed to support the Orion Crew Module and to support future Constellation missions. A MFE would be used from Earth sea level conditions to the vacuum of space. The current Space Shuttle configuration utilizes an ammonia boiler and flash evaporator system to achieve cooling at all altitudes. With the MFE system, both functions are combined into a single compact package with significant weight reduction and improved freeze-up protection. The heat exchanger core is designed so that radial flow of the evaporant provides increasing cross-sectional area to keep the back pressure low. Its multiple layer construction allows for efficient scale up to the desired heat rejection rate.

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.

Thermal characterization was performed on a vapor compression heat pump using a novel, hybrid two phase loop design. Previous work on this technology has demonstrated its ability to provide passive phase separation and flow control based on capillary action. This provides high quality vapor to the compressor without relying on gravity-based phase separation or other active devices. This paper describes the subsequent work done to characterize evaporator performance under various startup scenarios, tilt angles, and heat loads. The use of a thermal expansion valve as a method to regulate operation was investigated. The effect of past history of use on startup behavior was also studied. Testing under various tilt angles showed evaporator performance to be affected by both adverse and favorable tilts for the given compressor. And depending on the distribution of liquid in the system upon startup, markedly different performance can result for the same system settings and heat loads.

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.

During an ExtraVehicular Activity (EVA), both the heat generated by the astronaut's metabolism and that produced by the Portable Life Support System (PLSS) must be rejected to space. The heat sources include the heat of adsorption of metabolic CO2, the heat of condensation of water, the heat removed from the body by the liquid cooling garment and the load from the electrical components. Although the sublimator hardware to reject this load weighs only 1.58 kg (3.48 lbm), an additional 3.6 kg (8 lbm) of water are loaded into the unit, most of which is sublimated and lost to space, thus becoming the single largest expendable during an eight-hour EVA. Using a radiator to reject heat from the astronaut during an EVA can reduce the amount of expendable water consumed in the sublimator. Last year we reported on the design and initial operational assessment tests of a novel radiator designated the Radiator And Freeze Tolerant heat eXchanger (RAFT-X).

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 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.

Progress on the delivery of the Vehicle Cabin Atmosphere Monitor (VCAM) is reported. VCAM is an autonomous trace-species detector to be used aboard the International Space Station (ISS) for atmospheric analysis. The instrument is based on a low-mass, low-power miniature preconcentrator, gas chromatograph, and Paul ion trap mass spectrometer (PCGC/MS) capable of measuring volatile constituents in a space vehicle or planetary outpost at sub-ppm levels. VCAM detects and quantifies 40 target compounds at their 180-day Spacecraft Maximum Allowable Concentration (SMAC) levels. It is designed to operate autonomously, maintenance-free, with a self-contained carrier and calibration gas supplies sufficient for a one-year lifetime. Two flight units will be delivered for operation in the ISS EXPRESS rack.