Search Results

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.

Tilting pad bearings are designed by hydrodynamic principles and have been utilized in applications carrying shaft thrust or radial loads in many mechanisms for decades. The object of this paper is to derive the optimized pivoting positions in the radial and circumferential directions of tilting pad thrust and radial bearings and to calculate minimum fuel film thickness for a given running condition of velocity, temperature, viscosity, bearing geometry, and loading forces. The Reynolds equation derived on the tilting pad bearing fluid model is simplified into a one dimensional equation and applied in two dimensions to solve for the minimum fluid film thickness from pressure distribution in the load-carrying analysis.

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.

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.

The Portable Life Support System (PLSS) emergency oxygen system is being reexamined for the next generation of suits. These suits will be used for transit to Low Earth Orbit, the Moon and to Mars as well as on the surface of the Moon and Mars. Currently, the plan is that there will be two different sets of suits, but there is a strong desire for commonality between them for construction purposes. The main purpose of this paper is to evaluate what the emergency PLSS requirements are and how they might best be implemented. Options under consideration are enlarging the tanks on the PLSS, finding an alternate method of storage/delivery, or providing additional O2 from an external source. The system that shows the most promise is the cryogenic oxygen system with a composite dewar which uses a buddy system to split the necessary oxygen between two astronauts.

The Sensory Prognostics and Management Systems (SPMS) program sponsored by the Federal Aviation Administration and Boeing developed and evaluated designs to integrate advanced diagnostic and prognostic (i.e., Integrated Vehicle Health Management (IVHM) or Health Management (HM)) capabilities onto commercial airplanes. The objective of the program was to propose an advanced HM system appropriate for legacy and new aircraft and examine the technical requirements and their ramifications on the current infrastructure and regulatory guidance. The program approach was to determine the attractive and feasible HM applications, the technologies that are required to cost effectively implement these applications, the technical and certification challenges, and the system level and business consequences of such a system.

The Major Constituent Analyzer (MCA) is a mass spectrometer-based atmospheric monitoring instrument in the Laboratory Module of the International Space Station (ISS). The MCA is used for continuous environmental monitoring of 6 major gas constituents in the ISS atmosphere as well as safety-critical monitoring for special Environmental Control and Life Support (ECLS) operations such as Pre-Breathe in the Airlock for Extra-Vehicular Activities (EVAs) and oxygen re-pressurizations. For the latter, it is desirable to make most efficient use of consumables by transferring the maximum amount from O2 re-supply tanks on board the shuttle or Progress. The upper safety limit for O2 transfer is constrained by the MCA measurement error bands. A study was undertaken to tighten these error bands and afford NASA-Mission Operations Directorate (MOD) more operational flexibility.

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.

Carrier Injection Sensorless (CIS) rotor position estimation for electric machines depends on the rotor saliency “seen” from the stator terminals. Compared to permanent magnet and induction machines, wound field synchronous machine rotor saliency characteristics are more complex. At typical carrier frequencies subtransient characteristics dominate. Injecting positively rotating carrier voltages at the stator terminals of a machine having rotor saliency produces both positively and negatively rotating carrier currents. The impedances relating currents to injected carrier voltages are derived from the dq reference frame synchronous machine model. Analytical results are compared to simulation and limited test results for an aircraft starter/generator.

A High Heat Flux Demonstration Program has been initiated to investigate and demonstrate the performance of a number of candidate cooling technologies to address the need of dissipating the large thermal loads and high heat fluxes associated with Directed Energy Weapons (DEW) systems. The technologies selected for these investigations utilize both single-phase and two-phase cooling concepts. The single-phase devices investigated are based upon the concept of jet impingement with and without extended surface areas. The two-phase devices investigated extend the jet impingement concepts into the liquid-vapor phase change regime, as well as a device based upon vapor injection spray cooling technology. In addition, all devices must demonstrate scalability. For each device a unit cooling cell has been defined and greater surface area capability is to be achieved with the addition of adjacent cells without significantly affecting the performance of neighboring cells.

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.

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.

As next generation space suit concepts enable extravehicular activity (EVA) mission capability to extend beyond anything currently available today, revolutionary advances in life support technologies are required to achieve anticipated NASA mission profiles than may measure years in duration and require hundreds of sorties. Since most life support systems require power, increased mass and volume efficiency of the energy storage materials can have a dramatic impact on reducing the overall weight of next generation space suits. ITN Energy Systems, in collaboration with Hamilton Sundstrand and the NASA Johnson Space Center's EVA System's Team, is developing multifunctional fiber batteries to address these challenges. By depositing the battery on existing space suit materials, e.g. scrim fibers in the thermal micrometeoroid garment (TMG) layers, parasitic mass (inactive materials) is eliminated leading to effective energy densities ∼400 Wh/kg.

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.

In a crewed spacecraft environment, atmospheric carbon dioxide (CO2) and moisture control are crucial. Hamilton Sundstrand has developed a stable and efficient amine-based CO2 and water vapor sorbent, SA9T, that is well suited for use in a spacecraft environment. The sorbent is efficiently packaged in pressure-swing regenerable beds that are thermally linked to improve removal efficiency and minimize vehicle thermal loads. Flows are all controlled with a single spool valve. This technology has been baselined for the new Orion spacecraft. However, more data was needed on the operational characteristics of the package in a simulated spacecraft environment. A unit was therefore tested with simulated metabolic loads in a closed chamber at Johnson Space Center during the last third of 2006. Tests were run at a variety of cabin temperatures and with a range of operating conditions varying cycle time, vacuum pressure, air flow rate, and crew activity levels.

Under a NASA-sponsored technology development project, a multi-disciplinary team consisting of industry, academia, and government organizations led by Hamilton Sundstrand is developing an amine based humidity and carbon dioxide (CO2) removal process and prototype equipment for Vision for Space Exploration (VSE) applications. This system employs thermally linked amine sorbent beds operating as a pressure swing adsorption system, using the vacuum of space for regeneration. The prototype hardware was designed based on a two fault tolerant requirement, resulting in a single system that could handle the metabolic water and carbon dioxide load for a crew size of six. Two, full scale prototype hardware sets, consisting of a linear spool valve, actuator and amine sorbent canister, have been manufactured, tested, and subsequently delivered to NASA JSC. This paper presents the design configuration and the pre-delivery performance test results for the CAMRAS hardware.

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.

As next generation space suit concepts enable extravehicular activity (EVA) mission capability to extend beyond anything currently available today, revolutionary advances in life support technologies are required to achieve anticipated NASA mission profiles that may measure years in duration and require hundreds of sorties. Since most life support systems require power, increased mass and volume efficiency of the energy storage materials can have a dramatic impact on reducing the overall weight of next generation space suits. This paper details the development of a multifunctional fiber battery to address these needs.

The Orion Crew Exploration Vehicle (CEV) requires a smoke detector for the detection of particulate smoke products as part of the Fire Detection and Suppression (FDS) system. The smoke detector described in this paper is an adaptation of a mature commercial aircraft design for manned spaceflight. Changes made to the original design include upgrading the materials and electronics to space-qualified components, and modifying the mechanical design to withstand launch and landing loads. The results of laboratory characterization of the response of the new design to test particles are presented.