Search Results

Increased nickel concentrations in the IATCS coolant prompted a study of the corrosion rates of nickel-brazed heat exchangers in the system. The testing has shown that corrosion is occurring in a silicon-rich intermetallic phase in the braze filler of coldplates and heat exchangers as the result of a decrease in the coolant pH brought about by cabin carbon dioxide permeation through polymeric flexhoses. Similar corrosion is occurring in the EMU de-ionized water loop. Certain heat exchangers and coldplates have more silicon-rich phase because of their manufacturing method, and those units produce more nickel corrosion product. Silver biocide additions did not induce pitting corrosion at silver precipitate sites.

Under a NASA sponsored technology development activity, Hamilton Sundstrand has designed, fabricated, tested and delivered a prototype solid amine-based carbon dioxide (CO2) and water (H2O) vapor removal system sized for Extravehicular Activity (EVA) operation. The prototype system employs two alternating and thermally-linked solid amine sorbent beds to continuously remove CO2 and H2O vapor from a closed environment. While one sorbent bed is exposed to the vent loop to remove CO2 and water vapor, the other bed is exposed to a regeneration circuit, defined as either vacuum or an inert sweep gas stream. A linear spool valve, coupled directly to the amine canister assembly, is utilized to simultaneously divert the vent loop flow and regeneration circuit flow between the two sorbent beds.

Metabolically produced carbon dioxide (CO2) removal in spacesuit applications has traditionally been accomplished utilizing non-regenerative Lithium Hydroxide (LiOH) canisters. In recent years, regenerative Metal Oxide (MetOx) has been developed to replace the Extravehicular Mobility Unity (EMU) LiOH canister for extravehicular activity (EVA) missions in micro-gravity, however, MetOx may carry a significant weight burden for potential use in future Lunar or planetary EVA exploration missions. Additionally, both of these methods of CO2 removal have a finite capacity sized for the particular mission profile. Metabolically produced water vapor removal in spacesuits has historically been accomplished by a condensing heat exchanger within the ventilation process loop of the suit life support system.

An extremely compact heat exchanger is being developed which can boil cryogenic fluids with a liquid heat source at temperatures close to its freezing point. Freezing of the heat source fluid, e.g. water is precluded by the normal flow arrangement. Boiling and superheating of the cryogen occurs as the fluid approaches the heat source in a stack of bonded jet-array laminations. This heat exchanger technology is important in many applications where the storage of fluids at cryogenic temperatures offers substantial advantages in terms of system weight and volume. Often, as in several advanced portable life support system concepts, the advantages include the use of the cryogen as a heat sink in system thermal management. Realizing this benefit and safely conditioning the stored fluid for use requires effective heat transfer between the cryogen and a secondary heat transport fluid.

In future lunar and Mars exploration missions the ability to provide the crewmember navigation information will be critical. Exploration demands that Extravehicular Activity (EVA) astronauts be provided the capability to operate with greater autonomy in accomplishing complex EVA missions than has been the case previously. Robust crew information interfaces and navigation support integral to the EVA spacesuit system are expected to be minimum requirements. Navigation support must allow the EVA crew to determine their position relative to EVA target locations, satellite imagery and maps and assist them in walking or riding to the desired targets on the planetary surface. Together, these needs suggest a requirement for an integrated system that combines data and voice communications, a high performance visual display, and navigation support in a design that is compatible with spacesuit environmental and packaging restrictions and with unique EVA crew interface demands.

This paper addresses the issues of data reduction, visualization, clustering and classification for fault diagnosis and prognosis of the Liquid Cooling System (LCS) in an aircraft. LCS is a cooling system that consists of a left and a right loop, where each loop is composed of a variety of components including a heat exchanger, source control units, a compressor, and a pump. The LCS data and the fault correlation analysis used in the paper are provided by Hamilton Sundstrand (HS) - A United Technologies Company (UTC). This data set includes a variety of sensor measurements for system parameters including temperatures and pressures of different components, along with liquid levels and valve positions of the pumps and controllers. A graphical user interface (GUI) is developed in Matlab that facilitates extensive plotting of the parameters versus each other, and/or time to observe the trends in the data.

This paper addresses the issue of fault diagnosis in the heat exchanger of an aircraft Air Conditioning System (ACS). The heat exchanger cools the air by transferring the heat to the ram-air. Due to a variety of biological, mechanical and chemical reasons, the heat exchanger may experience fouling conditions that reduces the efficiency and could considerably affect the functionality of the ACS. Since, the access to the heat exchanger is limited and time consuming, it is preferable to implement an early fault diagnosis technique that would facilitate Condition Based Maintenance (CBM). The main contribution of the paper is pre-flight fault assessment of the heat exchanger using a combined model-based and data-driven approach of fault diagnosis. A Simulink model of the ACS, that has been designed and validated by an industry partner, has been used for generation of sensor data for various fouling conditions.

As part of NASA's initiative to develop an advanced portable life support system (PLSS), a baseline schematic has been chosen that includes gaseous oxygen in a closed circuit ventilation configuration. Supply oxygen enters the suit at the back of the helmet, passes over the astronaut's body, and is extracted at the astronaut's wrists and ankles through the liquid cooling and ventilation garment (LCVG). The extracted gases are then treated using a rapid cycling amine (RCA) system for carbon dioxide and water removal and activated carbon for trace gas removal before being mixed with makeup oxygen and reintroduced into the helmet. Thermal control is provided by a suit water membrane evaporator (SWME). As an extension of the original schematic development, NASA evaluated several Helmet Exhalation Capture System (HECS) configurations as alternatives to the baseline.

Following the Columbia accident, the EMUs (Extravehicular Mobility Units) onboard the ISS (International Space Station) went unused for an extended period of time. Upon startup, the units experienced a failure in the coolant systems. The failure resulted in a loss of EVA (Extravehicular Activity) capability from the US segment of the ISS. A failure investigation determined that chemical and biological contaminants and byproducts from the ISS Airlock Heat Exchanger, and the EMU itself, fouled the magnetically coupled pump in the EMU Transport Loop Fan/Pump Separator leading to a lack of coolant flow. Remediation hardware (the Airlock Coolant Loop Remediation water processing kit) and a process to periodically clean the EMU coolant loops on orbit were devised and implemented. The intent of this paper is to report on the successful implementation of the resultant hardware and process, and to highlight the go-forward plan.

Future spacesuit systems for the exploration of Mars will need to be much lighter than current designs, while at the same time reducing the consumption of water for crew cooling. One of the technology paths NASA has identified to achieve these objectives is the replacement of current high pressure oxygen storage technology in extravehicular activity (EVA) systems with cryogenic technology that can simultaneously reduce the mass of tankage required for oxygen storage and enable the use of the stored oxygen as a means of cooling the EVA astronaut. During the past year NASA has funded production of a prototype system demonstrating this capability in a design that will allow the cryogenic oxygen to be used in any attitude and gravity environment. This paper describes the design and manufacture of the prototype system. The potential significance and application of the system is also discussed.

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.

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.

Metabolic heat regenerated temperature swing adsorption (MTSA) technology is being developed for removal and rejection of carbon dioxide (CO2) and heat from a portable life support system (PLSS) to the Martian environment. Previously, hardware was built and tested to demonstrate using heat from simulated, dry ventilation loop gas to affect the temperature swing required to regenerate an adsorbent used for CO2 removal. New testing has been performed using a moist, simulated ventilation loop gas to demonstrate the effects of water condensing and freezing in the heat exchanger during adsorbent regeneration. Also, the impact of MTSA on PLSS design was evaluated by performing thermal balances assuming a specific PLSS architecture. Results using NASA's Extravehicular Activity System Sizing Analysis Tool (EVAS_SAT), a PLSS system evaluation tool, are presented.

Effective thermal and micrometeoroid protection as afforded by the Thermal Micrometeoroid Garment (TMG) is critical in achieving safe and efficient missions. It is also critical that the TMG does not increase torque or decreased range of motion which can cause crewmember discomfort, fatigue, and reduced efficiency. For future exploration missions, removable and replaceable TMGs will allow the use of different pressure garment protective covers and TMG configurations for launch, re-entry, 0-G Extra Vehicular Activity (EVA), and lunar surface EVA. A study was conducted with the goal of developing high Technology Readiness Level (TRL), scalable, interface design concepts for TMG systems. The affects of TMG segmentation on mobility and donning were assessed. Closure mechanisms were investigated and tested to determine their operability after exposure to lunar dust. A TMG configuration with the optimum number of segments and location of interfaces was selected for the Mark III spacesuit.