Results for “Venting valve” — Page 8

NASA NTRS2014

Thermal Vacuum Chamber Repressurization with Instrument Purging

At the conclusion of cryogenic vacuum testing of the James Webb Space Telescope Optical Telescope Element Integrated Science Instrument Module (JWST-OTIS) in NASA Johnson Space Center’s (JSCs) thermal vacuum (TV) Chamber A, contamination control (CC) engineers are postulating that chamber particulate material stirred up by the repressurization process may be kept from falling into the Integrated Science Instrument Module (ISIM) interior to some degree by activating instrument purge flows over some initial period before opening the chamber valves. This manuscript describes development of a series of models designed to describe this process. The models are strung together in tandem with a fictitious set of conditions to estimate overpressure evolution from which net outflow velocity behavior may be obtained. Creeping flow assumptions are then used to determine the maximum particle size that may be kept suspended above the ISIM aperture, keeping smaller particles from settling within the instrument module.

Repressurization↗

NASA NTRS2008

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively. The Mach Number distribution shows small yet distinct differences between the two cases such as locations of shocks/shock reflections and a slightly different impingement point on the wall of the diffuser from the expansion at the exit of the nozzle. Similarly the temperature distribution indicates different flow recirculation patterns in the test cell. Both cases capture all the essential flow phenomena such as the shock-boundary layer interaction, plume expansion, expansion of the first stage ejectors, mixing between the engine plume and the first stage ejector flow and pressurization due to the first stage ejectors. The final paper will discuss thermal loads on the walls of the diffuser and cooling mechanisms investigated.

Allgood, Daniel C.↗

NASA NTRS2010

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively.

Allgood, Daniel C.↗

NASA NTRS2014

Nuclear Thermal Rocket Element Environmental Simulator (NTREES) Upgrade Activities

Over the past year the Nuclear Thermal Rocket Element Environmental Simulator (NTREES) has been undergoing a significant upgrade beyond its initial configuration. The NTREES facility is designed to perform realistic non-nuclear testing of nuclear thermal rocket (NTR) fuel elements and fuel materials. Although the NTREES facility cannot mimic the neutron and gamma environment of an operating NTR, it can simulate the thermal hydraulic environment within an NTR fuel element to provide critical information on material performance and compatibility. The first phase of the upgrade activities which was completed in 2012 in part consisted of an extensive modification to the hydrogen system to permit computer controlled operations outside the building through the use of pneumatically operated variable position valves. This setup also allows the hydrogen flow rate to be increased to over 200 g/sec and reduced the operation complexity of the system. The second stage of modifications to NTREES which has just been completed expands the capabilities of the facility significantly. In particular, the previous 50 kW induction power supply has been replaced with a 1.2 MW unit which should allow more prototypical fuel element temperatures to be reached. The water cooling system was also upgraded to so as to be capable of removing 100% of the heat generated during. This new setup required that the NTREES vessel be raised onto a platform along with most of its associated gas and vent lines. In this arrangement, the induction heater and water systems are now located underneath the platform. In this new configuration, the 1.2 MW NTREES induction heater will be capable of testing fuel elements and fuel materials in flowing hydrogen at pressures up to 1000 psi at temperatures up to and beyond 3000 K and at near-prototypic reactor channel power densities. NTREES is also capable of testing potential fuel elements with a variety of propellants, including hydrogen with additives to inhibit corrosion of certain potential NTR fuel forms. Additional diagnostic upgrades included in the present NTREES set up include the addition of a gamma ray spectrometer located near the vent filter to detect uranium fuel particles exiting the fuel element in the propellant exhaust stream to provide additional information any material loss occurring during testing. Other aspects of the upgrade included reworking NTREES to reduce the operational complexity of the system despite the increased complexity of the induction heating system. To this end, many of the controls were consolidated on fewer panels. As part of this upgrade activity, the Safety Assessment (SA) and the Standard Operating Procedures (SOPs) for NTREES were extensively rewritten. The new 1.2 MW induction heater consists of three physical units consisting of a transformer, rectifier, and inverter. This multiunit arrangement facilitated increasing the flexibility of the induction heater by more easily allowing variable frequency operation. Frequency ranges between 20 and 60 kHz can be accommodated in the new induction heater allowing more representative power distributions to be generated within the test elements.

Emrich, William J., Jr.↗

NASA NTRS1990

Cryogenic On-Orbit Liquid Depot Storage, Acquisition, and Transfer Satellite (COLD-SAT)

The Cryogenic On-Orbit Liquid Depot Storage, Acquisition, and Transfer Satellite (COLD-SAT) will perform subcritical liquid hydrogen handling experiments under low gravity conditions to provide engineering data for future space transportation missions. Comprising the four Class 1 enabling experiments are tank press control, tank chilldown, tank no-vent fill, and liquid acquisition device fill/refill. The nine Class 2 enhancing experiments are tanker thermal performance, pressurization, low-gravity setting and outflow, liquid acquisition device performance, transfer line chilldown, outflow subcooling, low-gravity vented fill, fluid dumping, and advanced instrumentation. Consisting of an experiment module mated to a spacecraft bus, COLD-SAT will be placed in an initial 1300 km circular orbit by an Atlas commercial launch vehicle, and will perform experiments in a semi-autonomous mode for a period of up to six months. The three-axis controlled spacecraft bus provides electric power, control and data management, communications, and attitude control along with propulsive acceleration levels ranging from 10(exp -6) to 10(exp -4) g. It is desired to understand the effects that low acceleration levels might have on the heat and mass transfer processes involved in some of the experiments. The experiment module contains the three liquid hydrogen tanks, valves, pressurization and pumping equipment, and instrumentation. Within the highly insulated tanks are specialized fluid management equipment that might be used in future space transportation systems. At launch all the liquid hydrogen for the experiments is contained in the largest tank, which has helium-purged insulation to prevent cryo-pumping of air on the launch pad. The tank is loaded by the hydrogen tanking system used for the Centaur upper stage of the Atlas. After reaching orbit the two smaller tanks become receivers for fluid transfers, and when tanked, become the vessels for performing many of the experiments.

Schuster, John R.↗

NASA NTRS2014

Performing a Launch Depressurization Test on an Inflatable Space Habitat

In July, 2014 JPL's Environmental Test Laboratory successfully performed a launch depressurization test on an inflatable space habitat proposed to be installed on the International Space Station. The inflatable habitat is to be launched in the SpaceX Dragon Trunk. During the launch, the unpressurized Dragon Trunk will rapidly change from ground level atmospheric pressure to the vacuum of space. Since the inflatable habitat is tightly folded during launch with multiple layers of bladder, Kevlar fabric sections, and micro-meteoroid shielding, it was not possible to analyze or simulate how the residual air pockets would behave during the launch. If the inflatable habitat does not vent adequately and expands, it could rupture the payload bay of the launch vehicle. A launch depressurization test was chosen as the best way to qualify the inflatable habitat. When stowed, the inflatable habitat measured approximately 241 cm (95 inches) in diameter by 152 cm (60 inches) high and weighed close to 1361 kg (3,000 pounds). Two vacuum chambers connected by a large vacuum line were used to perform this test. The inflatable habitat was mounted in the smaller chamber, which was 396 cm (13 feet) in diameter and 1128 cm (37 feet) high. The larger chamber, which was 823 cm (27 feet) in diameter and 2,591 cm (85 feet) high, was rough pumped and used as a vacuum reservoir. A two stage axial type compressor and ten Stokes vacuum pumps were also used during the depressurization. Opening a butterfly valve on the vacuum line, at the smaller chamber, was manually controlled so that the smaller chamber's depressurization rate matched the launch pressure profile.

ETL↗

NASA NTRS2010

Trade Study for 9 kW Water Membrane Evaporator

Sublimators have been proposed and used in spacecraft for heat rejection. Sublimators are desirable heat rejection devices for short duration use because they can transfer large amounts of heat using little mass and are self-regulating devices. Sublimators reject heat into space by freezing water inside a porous substrate, allowing it to sublimate into vapor, and finally venting it into space. The state of the art thermal control system in orbiting spacecraft is a two loop, two fluid system. The external coolant loop typically uses a toxic single phase fluid that acquires heat from the spacecraft and rejects most of it via a radiator. The sublimator functions as a transient topper for orbiting spacecraft during day pass periods when radiator efficiency decreases. The sublimator interfaces with the internal loop through a built in heat exchanger. The internal loop fluid is non-toxic and is typically a propylene glycol and water solution with inhibitors to prevent corrosion with aluminum fins of the heat exchangers. Feedwater is supplied from a separate line to the sublimator to maintain temperature control of the cabin and vehicle hardware. Water membrane evaporators have been developed for spacecraft and spacesuits. They function similar to a sublimator but require a backpressure valve which could be actuated for this application with a simple fully open or fully closed modes. This technology would be applied to orbital thermal control (lunar or planetary). This paper details a trade study showing that evaporators would greatly reduce the consumable that is used, effectively wasted, by sublimators during start up and shut down during the topping phases of each orbit. State of the art for 9 kW sublimators reject about 870 W per kilogram of mass and 1150 W per liter of volume. If water with corrosion inhibitors is used the evaporators would be about 80% of the mass and volume of the equivalent system. The size and mass increases to about 110% if the internal fluid is 50% propylene glycol/50% water. The true benefit comes from the backpressure valve, that prevents the cyclical shutdown/startup loss of the sublimator and amounts to as much as 0.85 kg per orbit.

Bue, Grant C.↗

NASA NTRS2005

Reciprocating Feed System Development Status

The reciprocating feed system (RFS) is an alternative means of providing high pressure propellant flow at low cost and system mass, with high fail-operational reliability. The RFS functions by storing the liquid propellants in large, low-pressure tanks and then expelling each propellant through two or three small, high-pressure tanks. Each RFS tank is sequentially filled, pressurized, expelled, vented, and refilled so as to provide a constant, or variable, mass flow rate to the engine. This type of system is much lighter than a conventional pressure fed system in part due to the greatly reduced amount of inert tank weight. The delivered payload for an RFS is superior to that of conventional pressure fed systems for conditions of high total impulse and it is competitive with turbopump systems, up to approximately 2000 psi. An advanced version of the RFS uses autogenous pressurization and thrust augmentation to achieve higher performance. In this version, the pressurization gases are combusted in a small engine, thus making the pressurization system, in effect, part of the propulsion system. The RFS appears to be much less expensive than a turbopump system, due to reduced research and development cost and hardware cost, since it is basically composed of small high- pressure tanks, a pressurization system, and control valves. A major benefit is the high reliability fail-operational mode; in the event of a failure in one of the three tank-systems, it can operate on the two remaining tanks. Other benefits include variable pressure and flow rates, ease of engine restart in micro-gravity, and enhanced propellant acquisition and control under adverse acceleration conditions. We present a system mass analysis tool that accepts user inputs for various design and mission parameters and calculates such output values payload and vehicle weights for the conventional pressure fed system, the RFS, the Autogenous Pressurization Thrust Augmentation (APTA) RFS, and turbopump systems. Using this tool, a preliminary design of a representative crew exploration vehicle (CEV) has been considered. The design parameters selected for a representative system were modeled after the orbital maneuvering system (OMS) on the Shuttle Orbiter, with an increase of roughly a factor of ten in the delta- V capability and a greater thrust (30,000 lbs, vs. 12,000 lbs). Both storable and cryogenic propellants were considered. Results show that a RFS is a low mass alternative to conventional pressure fed systems, with a substantial increase in payload capability and that it is weight-competitive with turbopump systems at low engine pressure (a few hundred psi); at high engine pressures, the APTA RFS appears to offer the highest payload. We also present the status of the RFS test bed fabrication, assembly, and checkout. This test bed is designed to provide flow rates appropriate for engines in the roughly 10,000 to 30,000 lb thrust range.

Trewek, Mary↗

DOE OSTI2020

Tracking Radioactive Isotopes in HVAC and Application for Hot Cell Analyses - 20155

In recent years, the surge in the number of isotope production facilities under design has increased the need to analyze isotope migration in facility Heating, Ventilation, and Air Conditioning (HVAC) during normal operations and accident scenarios. These facilities are used to produce isotopes for medical, security, and industrial applications and as such are subject to license and regulatory requirements 10CFR20, 10CFR30, 10CFR50, and 10CFR70. GOTHIC, a general-purpose thermal-hydraulics software package, includes the ability to model isotopic tracers, radioactive decay and isotope migration as well as HEPA and charcoal filters for isotope retention. A GOTHIC model was developed by Zachry Nuclear Engineering (ZNE) to examine the effect of negative room pressure, HEPA filtration, and HVAC fluctuations on radiation areas, hot cells and gloveboxes. Radioactive tracers were used to simulate the concentration of spills within contaminated areas, track the migration of isotopes of interest, and determine the isotopic retention and buildup on facility HEPA filters. A variety of isotopes with concern to dose (e.g., Kr-85, Sr-90, I-131, etc.), including their decay and progeny, are included in the analysis. Negative pressures are maintained in the regions of interest by a representative central HVAC system equipped with a volumetric fan that exhausts to the environment after a series of isolation valves and HEPA filters. GOTHIC is an industry trusted tool for providing engineering solutions for a variety of applications, including fission product tracking, aerosol and particulate transport and ventilation assessments. The software provides an integrated analysis environment that includes a graphical user interface (GUI) for constructing analysis models, a numerical solver that includes parallel processing capabilities and a post-processor for evaluating simulation results. It solves the conservation equations for mass, momentum and energy for multicomponent, multi-phase flow in lumped parameter and multi-dimensional geometries (1, 2, or full 3D), including the effects of turbulence, diffusion and buoyancy. It has been developed and maintained under a Quality Assurance program in compliance with the requirements of 10CFR50 Appendix B and applicable portions of ASME NQA-1 since 1995. GOTHIC has been used for assessing both forced and natural convection conditions for a wide range of applications, including: - Tracking concentration of hazardous gases and chemicals for habitability and safety assessments - Determining ventilation and filtration requirements and optimizing location and arrangement of these systems - Room heat-up, including diverse and Flexible coping strategies for Extended Loss of AC Power (FLEX/ELAP) - Equipment Qualification (EQ) A distinctive feature of GOTHIC is the ability to track many different fields/substances in a simulation, including user defined tracer elements, in the liquid, vapor and droplet fields as well as surfaces and filters. This capability allows GOTHIC to model fission product transport and release or the removal of particulates or harmful toxins from exhaust gases using a spray scrubber or other types of filtration systems. GOTHIC also includes models for engineered equipment, such as fans, filters, charcoal filters, dryers/demisters, dampers, etc. The aerosols and other filtered material are removed or accumulated in these components. The range of aerosol and radiological applications that GOTHIC has been used for includes: - Source Term: Primary Coolant (Equilibrium) Activity; Non-Water Coolant Source/Leakage. - Conditions for Iodine Re-evolution: Sump/Suppression Pool Conditions and pH; - RWST Conditions and pH. - Isotope Removal Mechanisms: Containment Sprayed and Unsprayed Region Mixing; Charcoal Filter Heating due to Iodine decay. - Radionuclide Transport and Decay: Post-LOCA Release in containment; Transport between connected Compartments and vent systems; Groundwater transport of radionuclides. - Non-Newtonian Fluid modeling for sludge, waste tanks, etc. ADAMS ML071581053 (titled 'Best Practice Guidelines for the use of CFD in Nuclear Reactor Safety Applications') poses guidelines for applying single phase CFD codes in nuclear reactor safety problems and GOTHIC is listed as a 'tool for 3D flows' and 'dispersal and deposition of radionuclides.' The Nuclear Quality Assurance (NQA) pedigree of GOTHIC is an important aspect for applications in the nuclear industry. The fundamental tracer models (convective transport, molecular and turbulent diffusion, removal mechanisms, etc.) have been verified using analytical solutions and validated against applicable separate effects tests. Also, GOTHIC has been benchmarked to many integrated effects tests, including Phebus FP (Fission Product). GOTHIC gives good agreement for the buildup and decay of fission products in Phebus Test 3. The model developed by ZNE demonstrates GOTHIC's applicability and acceptability for use in analyzing the migration and retention of radioactive isotopes and their progeny in normal operation and accident scenario analyses for isotope production facilities, hot cells, and gloveboxes. The tracer activities calculated by GOTHIC can then be used in downstream radiation transport and shielding codes like RADTRAD-NAI{sup C}, MCNP{sup R}, and MicroShield{sup R} to determine on-site and of-site doses. (authors)

12 MANAGEMENT OF RADIOACTIVE AND NON-RADIOACTIVE W↗

NASA NTRS2010

Gas Flow Detection System

This system provides a portable means to detect gas flow through a thin-walled tube without breaking into the tubing system. The flow detection system was specifically designed to detect flow through two parallel branches of a manifold with only one inlet and outlet, and is a means for verifying a space shuttle program requirement that saves time and reduces the risk of flight hardware damage compared to the current means of requirement verification. The prototype Purge Vent and Drain Window Cavity Conditioning System (PVD WCCS) Flow Detection System consists of a heater and a temperature-sensing thermistor attached to a piece of Velcro to be attached to each branch of a WCCS manifold for the duration of the requirement verification test. The heaters and thermistors are connected to a shielded cable and then to an electronics enclosure, which contains the power supplies, relays, and circuit board to provide power, signal conditioning, and control. The electronics enclosure is then connected to a commercial data acquisition box to provide analog to digital conversion as well as digital control. This data acquisition box is then connected to a commercial laptop running a custom application created using National Instruments LabVIEW. The operation of the PVD WCCS Flow Detection System consists of first attaching a heater/thermistor assembly to each of the two branches of one manifold while there is no flow through the manifold. Next, the software application running on the laptop is used to turn on the heaters and to monitor the manifold branch temperatures. When the system has reached thermal equilibrium, the software application s graphical user interface (GUI) will indicate that the branch temperatures are stable. The operator can then physically open the flow control valve to initiate the test flow of gaseous nitrogen (GN2) through the manifold. Next, the software user interface will be monitored for stable temperature indications when the system is again at thermal equilibrium with the test flow of GN2. The temperature drop of each branch from its "no flow" stable temperature peak to its stable "with flow" temperature will allow the operator to determine whether a minimum level of flow exists. An alternative operation has the operator turning on the software only long enough to record the ambient temperature of the tubing before turning on the heaters and initiating GN2 flow. The stable temperature of the heated tubing with GN2 flow is then compared with the ambient tubing temperature to determine if flow is present in each branch. To help quantify the level of flow in the manifolds, each branch will be bench calibrated to establish its thermal properties using the flow detection system and different flow rates. These calibration values can then be incorporated into the software application to provide more detailed flow rate information.

Moss, Thomas↗

NASA NTRS2026

Dynamic Regulation of Sub-Atmospheric Pressure for Constant and Cyclic Gas Loads During Testing of Spacesuit Components

New iterations of the various subsystems within the spacesuit will benefit from a new sorbent technology. For example, sorbents used in the Trace Contaminant Control (TCC) and the Rapid Cycle Amine (RCA) systems within the Exploration Portable Life Support System (xPLSS), part of the Extra-Vehicular Mobility Unit (xEMU). For proper validation, the sub-atmospheric pressure needs to be maintained within a simulated 2 ft 3 spacesuit volume. The traditional approach, which involves placing the entire vent loop within a hypobaric chamber, is not practical for the widespread testing of components and prototypes, as these specialized chambers are costly and not widely available. A two-stage regulator based sub-atmospheric pressure system is presented in this work. The method regulates the pressure based on pressure differentials between the pressure regulation pump, the test system, and the ambient environment. The performance of this system is demonstrated using data collected during a 150+ hour non-regenerative TCC sorbent evaluation, several 24-hour regenerative TCC sorbent evaluations with different regeneration cycles, and an 8-hour RCA sorbent evaluation, all under xEMU operating conditions. For regulation, the inlet regulator was set to propagate a small leak to increase the stability of the system as the gas load from the testing system changes. It was also found that controlling the input flow rate that replaces the material lost during regeneration is critical for ensuring the stability of the system. This system provided excellent pressure regulation, without adjustments, for constant loads during shorter time evaluations (hours), while longer time evaluations (days) are easily obtainable with periodic regulator adjustments. A second method is under development to address longer-term stability and to automate pressure regulation. It will incorporate flow and pressure measurements to dynamically control a set of electronically controlled proportional valves for adjusting the pumping speed and air bleed supply.

Nicholas F Materer↗

NASA NTRS2009

Pressure-Energized Seal Rings to Better Withstand Flows

Pressure-energized seal rings intended to withstand flows better than do conventional pressure-energized seal rings have been conceived. The concept applies, more specifically, to seal rings used on some valve stems, pistons, and the like. A conventional pressure-energized seal ring generally has a U-shaped cross section and consists of an elastomer or other suitable polymer with an embedded metal energizing spring (see Figure 1). The working fluid from the high-pressure side that one seeks to seal is allowed into the U-shaped cavity, so that the pressure pushes the sides of the seal ring tighter against the gland and body sealing surfaces, thereby increasing the degree of sealing. Unfortunately, when the seal ring is exposed to flow of the working fluid, under some conditions, the flow grabs the lip of the U-shaped cross section and ejects or deforms the seal ring so that, thereafter, a proper seal is not obtained. Figure 2 depicts one of several alternative seal rings according to the present concept. One element of the concept is to modify the U-shaped cross section from that of the corresponding conventional seal ring to eliminate the exposed lip and prevent entry of the working fluid into the U-shaped cavity. Unlike in the conventional seal, pressurized fluid would not push the seal ring directly against the both gland and body sealing surfaces. Instead, the pressure would directly push the seal ring against a gland sealing surface only. In so doing, the pressure would squash the seal ring into a smaller volume bounded by the gland and body sealing surfaces, and would thereby indirectly press the seal ring more tightly against the body sealing surface. To enhance the desired squashing deformation, a spring having an approximately parallelogram cross section would be embedded in the modified U-shaped cavity. As the pressure pushed two corners of the approximate parallelogram closer together along the axis of the seal ring, the other two corners of the approximate parallelogram would be pushed farther apart along a radius of the ring, thereby causing the polymeric ring material to push radially harder against the body sealing surface. From the radially innermost corner of the approximate parallelogram, the spring material would extend radially, then axially into recesses in the seal gland. These extensions would help to restrain the seal ring against ejection. A seat retainer would hold the sealing ring in the gland and form a mechanical compression seal to prevent or at least reduce leakage of pressurized fluid into the cavity behind the seal. However, because there would likely be a little leakage, the cavity behind the seal should be vented to the low pressure side to prevent buildup of pressure in the cavity over time; otherwise, the built-up pressure could cause ejection of the seal ring when the pressure on the high-pressure side was reduced. Polymeric seal-ring materials may not be able to withstand working conditions in applications that involve abrasive and/or hot working fluids. For such applications, all-metal seal rings may be preferred. The bottom part of Figure 2 shows one example of an alternative gland configuration with an all-metal seal ring.

Farner, Bruce↗