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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Variable Emissivity Through MEMS Technology

This paper discusses a new technology for variable emissivity (vari-e) radiator surfaces, which has significant advantages over traditional radiators and promises an alternative design technique for future spacecraft thermal control systems. All spacecraft rely on radiative surfaces to dissipate waste heat. These radiators have special coatings, typically with a low solar absorptivity and a high infrared-red emissivity, that are intended to optimize performance under the expected heat load and thermal sink environment. The dynamics of the heat loads and thermal environment make it a challenge to properly size the radiator and often require some means of regulating the heat rejection rate of the radiators in order to achieve proper thermal balance. Specialized thermal control coatings, which can passively or actively adjust their emissivity offer an attractive solution to these design challenges. Such systems would allow intelligent control of the rate of heat loss from a radiator in response to heat load and thermal environmental variations. Intelligent thermal control through variable emissivity systems is well suited for nano and pico spacecraft applications where large thermal fluctuations are expected due to the small thermal mass and limited electric resources. Presently there are three different types of vari-e technologies under development: Micro ElectroMechanical Systems (MEMS) louvers, Electrochromic devices, and Electrophoretic devices. This paper will describe several prototypes of micromachined (MEMS) louvers and experimental results for the emissivity variations measured on theses prototypes. It will further discuss possible actuation mechanisms and space reliability aspects for different designs. Finally, for comparison parametric evaluations of the thermal performances of the new vari-e technology and standard thermal control systems are presented in this paper.

Darrin, Ann Garrison↗
NASA NTRS2011

Design Considerations for Fusible Heat Sink

Traditionally radiator designs are based off a passive or flow through design depending on vehicle requirements. For cyclical heat loads, a novel idea of combining a full flow through radiator to a phase change material is currently being investigated. The flow through radiator can be designed for an average heat load while the phase change material can be used as a source of supplemental heat rejections when vehicle heat loads go above the average load. Furthermore, by using water as the phase change material, harmful radiation protection can be provided to the crew. This paper discusses numerous trades conducted to understand the most optimal fusible heat sink design for a particular heat load. Trades include configuration concepts, amount of phase change needed for supplemental heat rejection, and the form of interstitial material needed for optimal performance. These trades were used to culminate to a fusible heat sink design. The paper will discuss design parameters taken into account to develop an engineering development unit.

Cognata, Thomas J.↗
NASA NTRS2007

Lunar Return Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Wall Structures

Thermostructural analysis was performed on generic crew exploration vehicle (GCEV) heat shielded wall structures subjected to reentry heating rates based on five potential lunar return reentry trajectories. The GCEV windward outer wall is fabricated with a graphite/epoxy composite honeycomb sandwich panel and the inner wall with an aluminum honeycomb sandwich panel. The outer wall is protected with an ablative Avcoat-5026-39H/CG thermal protection system (TPS). A virtual ablation method (a graphical approximation) developed earlier was further extended, and was used to estimate the ablation periods, ablation heat loads, and the TPS recession layer depths. It was found that up to 83 95 percent of the total reentry heat load was dissipated in the TPS ablation process, leaving a small amount (3-15 percent) of the remaining total reentry heat load to heat the virgin TPS and maintain the TPS surface at the ablation temperature, 1,200 F. The GCEV stagnation point TPS recession layer depths were estimated to be in the range of 0.280-0.910 in, and the allowable minimum stagnation point TPS thicknesses that could maintain the substructural composite sandwich wall at the limit temperature of 300 F were found to be in the range of 0.767-1.538 in. Based on results from the present analyses, the lunar return abort ballistic reentry was found to be quite attractive because it required less TPS weight than the lunar return direct, the lunar return skipping, or the low Earth orbit guided reentry, and only 11.6 percent more TPS weight than the low Earth orbit ballistic reentry that will encounter a considerable weight penalty to obtain the Earth orbit. The analysis also showed that the TPS weight required for the lunar return skipping reentry was much more than the TPS weight necessary for any of the other reentry trajectories considered.

Ko, William L.↗
NASA NTRS2003

Progress Toward a Compact 0.05 K Magnet Refrigerator Operating from 10 K

Much of the most interesting information regarding our universe is hidden in the sub-millimeter, infrared, and x-rays bands of the spectrum, to which our atmosphere is largely opaque. Thus, missions exploring these bands are a very important part of NASA s Space Science program. Coincidentally, the most sensitive detectors in these spectral regions operate at extremely low temperatures, typically 0.05 - 0.10 K. Generally these temperatures will be achieved using magnetic refrigerators, also know as Adiabatic Demagnetization Refrigerators, or ADRs. Current ADRs, such as the one used in the XRS-II instrument on the Astro-E2 satellite, use a single-stage to cool detectors from 1.3 K to 0.06 K. The ADR is designed so that it can absorb the heat on the detector stage for at least 24 hours before it must stop, warm up to the helium bath temperature (1.3 K), and dump the accumulated heat. Future detector arrays will be much larger and will have higher heat dissipation. Furthermore, future missions will use mechanical cryocoolers to provide upper stage cooling, but they can only reach 4 - 10 K. Trying to scale heavy (-15 kg) single stage ADRs up to the higher heat loads and higher heat rejection temperatures required leads to unacceptably large systems. The GSFC Cryogenics Branch has developed the Continuous ADR (CADR) to solve this problem. The CADR consists of a series of ADR stages that sequentially pass heat from the load up to the high temperature heat sink. The stage connected to the load remains at a constant temperature. The continuous stage effectively decouples detector operation from ADR operation, allowing the ADR stages to be cycled much more rapidly. Rapid cycling leads to higher cooling power density. The cascading, multistage arrangement allows the magnetic refrigerant of each stage to be optimized for its own temperature swing. In the past year, we have made good progress toward a 0.05 to 10K system. A four-stage system that operates from 4.2 K was demonstrated. Magnetic shielding was added to eliminate inter-stage coupling. Improvements were made to superconducting and passive gas-gap heat switches. A second type of passive gas gap switch, one meant for use at higher temperature, was demonstrated. The presentation will focus primarily on these recent design improvements, and on the challenges that remain on the progress toward a system that will operate from 10 K or higher.

Canavan, Edgar↗
NASA NTRSDate not supplied

Control System Development for A Zero Boil-Off Hydrogen Storage Demonstration With Two-Stage Active Cooling

A NASA team is designing and building a test article to demonstrate the long duration storage of liquid hydrogen via active cooling (cryocoolers) while implementing a two-stage cooling approach. This activity is one of a large portfolio of NASA’s Space Technology Mission Directorate funded activities which focuses on the technology maturation needed for long duration storage of cryogenic liquid in-space. The current state-of-the-art for liquid hydrogen storage on-orbit is on the order of hours while NASA’s planned missions require storage for months, or even years. To enable such a long duration, Zero Boil-Off conditions must be achieved which requires passive technologies to minimize environmental heat loads, but also active cooling to intercept and reject the remaining heat to deep space. The implementation of active cooling results in a significant amount of dry mass added to the vehicle. Utilizing a Two-Stage Cooling approach, analysis and testing to date indicates the Zero Boil-Off of liquid hydrogen can be achieved with less mass and electrical power relative to a single-stage cooling approach where only one cryocooler is used. This activity will demonstrate the fully integrated suite of technologies needed to enable the Two-Stage Cooling approach which includes Multi-Layer Insulation blankets, Low Conductivity Structures, a Tube-On-Tank Heat Exchanger, Tube-On-Shield Heat Exchanger and a Two-Stage Cryogenerator (both 90 Kelvin and 20 Kelvin) with each stage having an independently controlled circulation loop. The bulk of the heat load is intercepted by the 90 Kelvin loop via a thin foil heat exchanger internal to the insulation blankets (also known as a Broad Area Cooling Shield), with the remainder removed at cryofluid temperature (20 Kelvin) by tubes directly welded to the outer tank surface. In this paper, we examine the challenges of controlling the system, and describe the development of control algorithms and software for future testing with liquid hydrogen. The interactions between the cryofans which control circulation loop mass flow rates, electrical heaters which vary the cryocooler lift to simulate the operation of variable-lift flight units, tank heaters which can vary the overall heat load, and the effects of changes in circulation loop pressure as temperatures change, can be complex and must be well understood to maintain steady-state propellant conditions and achieve Zero Boil-Off. Parallel PID loops and watchdog programs implemented on the user interface system will help bring the systems to steady state operation and keep them at selected operating points within acceptable error, while avoiding runaway feedback loops.

cryogenics↗
DOE OSTI2024

Design Heating and Cooling Load Calculation Versus Building Load Simulation for Cold Climate Heat Pumps: Understanding the "Gap"

This report explores the differences between Manual J-equivalent block load calculations and building HVAC energy simulation results using EnergyPlus™ calculations when designing cold climate heat pump systems for residential use. This study will help HVAC researchers and advanced designers understand the impacts of oversizing heat pumps on home energy use.

32 ENERGY CONSERVATION, CONSUMPTION, AND UTILIZATI↗
DOE OSTI2023

Field Validation of Air-Source Heat Pumps for Cold Climates

Heating energy is the largest end-use for U.S. residential buildings accounting for approximately one-third of residential building energy consumption (EIA 2021). Historically, air-source heat pumps have been limited to temperate climates because of subpar performance at extremely cold outdoor air temperatures. However, recent advances to cold-climate air-source heat pump technology, which typically rely on inverter-driven, variable-speed compressors and variable-speed fans, have significantly improved low-temperature heat pump performance enabling the technology to save energy for many homes in cold climates. The primary objective of this project was to measure in-field performance of centrally ducted, variable-capacity air-source heat pumps in cold climates to validate performance and develop field-based performance maps. The project focused on quantifying heat pump performance at cold temperatures. The sites identified for the study were primarily located in the Northwest United States since homes in the region tend to have all-electric space heating systems and high-efficiency heat pumps have been incentivized in the region for several years. NREL partnered with Ecotope, Inc., a small energy consulting firm located in Seattle, WA, for site recruitment, monitoring equipment installation, data quality management. All the sites included in the study had previously installed a high-efficiency, central heat pump system. One site was in a Denver, CO suburb, which was the only dual fuel heat pump in the study. We used airside and power measurements, collected at 5-second intervals, to quantify heat pump capacity, coefficient of performance (COP), and auxiliary heater energy consumption. We developed algorithms to automatically determine the heat pump operating mode including defrost and auxiliary heating operation. A whole-house thermal and duct audit was completed during the initial site visit to estimate winter heating loads and assess heat pump sizing. Whole-home heating design loads were calculated at ASHRAE 99% design temperatures and compared to manufacturer-reported maximum capacities to assess the heat pump sizing at each site.

32 ENERGY CONSERVATION, CONSUMPTION, AND UTILIZATI↗
NASA NTRS2004

Freeze Tolerant Radiator for an Advanced EMU

During an Extravehicular Activity (EVA), the astronaut s metabolic heat and the heat produced by the Portable Life Support Unit (PLSS) must be rejected. This heat load is currently rejected by a sublimator, which vents up to eight pounds of water each EVA. However, for advanced space missions of the future, water venting to space needs to be minimized because resupply impacts from earth will be prohibitive. If this heat load could be radiated to space from the PLSS, which has enough surface area to radiate most of the heat, the amount of water now vented could be greatly reduced. Unfortunately, a radiator rejects heat at a relatively constant rate, but the astronauts generate a variable heat load depending on how hard they are working. Without a way to vary the heat removal rate, the astronaut would experience cold discomfort or even frostbite. A proven method allowing a radiator to be turned-down is to sequentially allow tubes that carry the heat transfer fluid to the radiator to freeze. A drawback of current freezable radiators using this method is that they are far to heavy for use on a PLSS, because they use heavy construction to prevent the tubes from bursting as they freeze and thaw. This creates the need for a large radiator to reject most of the heat but with a lightweight tube that doesn t burst as it freezes and thaws. The new freezable radiator for the Extravehicular Mobility Unit (EMU) has features to accommodate the expansion of the radiator fluid when it freezes, and still have the high tube to fin conductance needed to minimize the number and weight of the tubes. Radiator fluid candidates are water and a propylene glycol-water mixture. This design maintains all materials within their elastic limits so that large volume changes can be achieved without breaking the tube. This concept couples this elastic expansion with an extremely lightweight, extremely high conductivity carbon fiber fin that can carry the heat needed to thaw a frozen tube. By using most of the exposed surface area of the PLSS as a radiator, the system can reject about 75% of the highest heat load, and reduce the loss of water through sublimation by a factor of four. The proposed radiator and a small water tank can be no heavier than the current system.

Copeland, Robert J.↗
NASA NTRS2021

Demonstration of Multilayer Insulation, Vapor Cooling of Structure, and Mass Gauging for Large Scale Upper Stages: Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER) Final Report

Testing was completed on the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER) between August 2019 and January 2020. SHIIVER was designed to be a test bed for the scaling of cryogenic fluid management technologies as applied to large upper stages and long duration in-space stages. The hardware consists of a 4 meter diameter stainless steel tank, structural skirts supporting the tank in the aft direction, and an aluminum forward skirt with vapor cooling channels bolted to it. The initial testing of the SHIIVER hardware was with liquid hydrogen and liquid nitrogen (as a substitute fluid for liquid oxygen and liquid methane) and sought to demonstrate the use of boil-off vapor to intercept heat on a structural skirt, multilayer insulation (MLI) on the tank domes, and the radio frequency mass gauge (RFMG). Testing was completed in four stages: a baseline thermal vacuum test prior to installation of the MLI, a thermal vacuum test after the MLI installation, a reverberant acoustic test, and a subsequent thermal vacuum test to verify that no damage occurred during the reverberant acoustic testing. Each thermal vacuum test with chamber wall at ambient temperature and vacuum level in the 10-6 torr range was conducted continuously between approximately 90% full and 25% full. Test results showed that the vapor cooling reduced the heat load to the tank by approximately 10%, but the boil-off by less than 3% at 50% full with no MLI on the domes. It reduced the heat load to the tank by approximately 10%, but the boil-off was essentially unchanged at 50% full whether or not vapor cooling was operational with MLI on the domes. The MLI reduced the heat load to the tank by approximately 40% at all fill levels, but the boil-off by approximately 25% at 90% full and 45% below 65% full. The RFMG performed well over all fill ranges, and several RF tank modes were used to gauge the mass of fluid in the tank. SHIIVER was then exposed to an acoustic environment of 147 dB OASPL (overall sound pressure level) in a reverberant chamber. The acoustic environment and profile envelopes the upper stage internal acoustic level of several different modern launch vehicles. No structural or thermal performance changes were observed as a result of acoustic testing. Final thermal vacuum testing after the acoustic testing showed no degradation to the MLI due to the acoustic environment as measured via system heat loads.

SHIIVER↗
NASA NTRS2014

Tank Applied Testing of Load-Bearing Multilayer Insulation (LB-MLI)

The development of long duration orbital cryogenic storage systems will require the reduction of heat loads into the storage tank. In the case of liquid hydrogen, complete elimination of the heat load at 20 K is currently impractical due to the limitations in lift available on flight cryocoolers. In order to reduce the heat load, without having to remove heat at 20 K, the concept of Reduced Boil-Off uses cooled shields within the insulation system at approximately 90 K. The development of Load-Bearing Multilayer Insulation (LB-MLI) allowed the 90 K shield with tubing and cryocooler attachments to be suspended within the MLI and still be structurally stable. Coupon testing, both thermal and structural was performed to verify that the LB-MLI should work at the tank applied level. Then tank applied thermal and structural (acoustic) testing was performed to demonstrate the functionality of the LB-MLI as a structural insulation system. The LB-MLI showed no degradation of thermal performance due to the acoustic testing and showed excellent thermal performance when integrated with a 90 K class cryocooler on a liquid hydrogen tank.

Thermal Insulation↗
NASA NTRS2014

Tank Applied Testing of Load-Bearing Multilayer Insulation (LB-MLI)

The development of long duration orbital cryogenic storage systems will require the reduction of heat loads into the storage tank. In the case of liquid hydrogen, complete elimination of the heat load at 20 K is currently impractical due to the limitations in lift available on flight cryocoolers. In order to reduce the heat load, without having to remove heat at 20 K, the concept of Reduced Boil-Off uses cooled shields within the insulation system at approximately 90 K. The development of Load-Bearing Multilayer Insulation (LB-MLI) allowed the 90 K shield with tubing and cryocooler attachments to be suspended within the MLI and still be structurally stable. Coupon testing both thermally and structurally were performed to verify that the LB-MLI should work at the tank applied level. Then tank applied thermal and structural (acoustic) testing was performed to demonstrate the functionality of the LB-MLI as a structural insulation system. The LB-MLI showed no degradation of thermal performance due to the acoustic testing and showed excellent thermal performance when integrated with a 90 K class cryocooler on a liquid hydrogen tank.

Space Environment Simulation↗
DOE OSTI2022

Thermal analysis of RHIC arc dipole magnet cold mass with EIC beam screen

The existing RHIC storage rings – including their superconducting magnet arcs – will be used for the hadron storage ring of the Electron-Ion Collider (EIC). The vacuum chamber of these magnets was not designed for the EIC hadron beams, with shorter bunches and of higher average current than the RHIC beams. With the current stainless steel beam pipe, the resistive-wall (RW) heating will exceed the dynamic heat load budget. Limiting the RW heating is important to prevent the superconducting magnets from quenching and to maintain a low screen temperature necessary to impede the rise of RW heating (higher resistance at higher temperature) as well as to achieve desired ultra-high vacuum. In addition, simulations predict the formation of electron cloud which would further contribute to the dynamic heat load and could compromise the quality and stability of the beam. To reduce the resistive-wall heating and suppress electron cloud, a beam screen will be installed in the vacuum chamber of the RHIC SC magnets. The screen will have a high RRR copper layer at its inner face – useful to reduce the resistive-wall impedance thanks to its high conductivity especially at cryogenic temperatures – and will be coated with a thin layer of amorphous carbon, a material with low secondary electron yield to suppress the formation of electron clouds. The baseline solution envisages a screen that will be cooled by thermal contact to the 4.55 K beam pipe. Detailed thermal analysis have been conducted in ANSYS 2020 for an arc dipole cold mass equipped with a beam screen in order to study the feasibility of a passively-cooled screen and guide its design. Temperature-dependent thermal conductivity properties of all materials in the operating (cryogenic) temperature range are considered. Suitable assumptions and simplifications are made to model the magnet coil and calculate its homogenized thermal conductivity. Sensitivity studies with respect to layer thicknesses and area of contact are carried out and results are presented.

43 PARTICLE ACCELERATORS↗
NASA NTRS1999

Flight Testing of a Cryogenic Capillary Pumped Loop

Future space-based cryogenic systems will require enhanced integration flexibility, lower weight reduced parasitic penalties, better vibration isolation, and a variety of other improvements to meet performance goals. Additionally, there is an increasing need to locate cooling sources remotely from cooled components. In the past flexible conductive links were used and worked well in most cases. However, as the transport lengths increase, conductive couplings become heavier and less effective, and must be replaced by higher performance systems. One available option, which can meet many of these future requirements, is the cryogenic capillary pumped loop (CCPL). The development of the CCPL technology started in 1992, following the success of the room temperature CPLS. The extrapolation of CCPL technology to cryogenic temperatures offers many performance benefits, which are not currently within the reach of traditional heat pipes or conductive links. Specific advantages of the CCPL technology pertaining to cryocooler integration include: (1) greater capillary pumping pressure for improved ground testability; (2) improved mechanical isolation; (3) faster diode shutdown and lower reverse heat leaks; (4) tighter control of detector temperature; (5) variable or fixed conductance operation; and (6) ease of integration due to their flexibility. The applications of CCPL technology are numerous. Military and commercial applications include surveillance satellites, earth observing satellites, deep space observation systems, medical devices, and many other cryogenic systems. Over the past few years, several breadboard and prototype CCPLs have been built and ground tested. A prototype CCPL has demonstrated successful operation between 80K and 110K with heat loads between O.5W and 12W using nitrogen as the working fluid, and 35K and 40K with head loads of 0.25W to 3.5W using neon. In order to verify CCPL performance in a microgravity environment, a flight unit, CCPL-5, was tested onboard the Space Shuttle STS-95 in October of 1998 as part of the CRYOTSU Flight Experiment. This flight was the first in-space demonstration of the CCPL. The CCPL-5 utilized nitrogen as the working fluid and operated between 75K and 110K. Flight results indicated excellent performance of the CCPL-5 under zero-G environment The CCPL could start from a supercritical condition in all tests, and the loop operating temperature could be tightly controlled regardless of changes in the heat load and/or the sink temperature. In addition, the loop demonstrated successful operation with a heat load of 0.5W as well as with parasitic heat loads alone. There were no noticeable differences between zero-G and one-G operation.

Ku, Jentung↗
DOE OSTI2022

Electron Cloud Simulations for the Electron-Ion Collider in Brookhaven National Laboratory

For high-intensity circular accelerators and storage rings, if the secondary electron emission yield (SEY) of the vacuum chamber surfaces is high, an EC (Electron Cloud) could build up with the passage of the circulating beam. The presence of EC can strongly affect the beam quality, such as transverse instabilities, transverse emittance growth, and beam loss. Furthermore, the heat load from EC can exceed the available cryogenic capability. For the RHIC superconducting (SC) arc magnets, to be used for the hadron storage ring of the Electron-Ion Collider (EIC), the dynamic heat load budget is 0.5 W/m to the 4.5 K stainless steel beam pipe. This can limit the maximum beam bunches or intensity of the EIC, hence diminishing the luminosity provided by the EIC. The EC has affected the beam instability and significantly contributed to cryogenic heat load in the Large Hadron Collider (LHC). Positron storage rings for which ECs have been a critical factor in the design and performance include KEKB in Japan and EC buildup remains one of the concerns for future high-intensity accelerators design. EC considerations have driven the SuperKEKB collider design and the positron damping ring for the proposed International Linear Collider (ILC). The LHC luminosity upgrade is contingent on reducing the bunch spacing to 25 ns; at this bunch spacing, severe EC buildup has been observed. The success of the upgrade is likely contingent on limiting EC buildup. To study the EC heat load, we did some EC simulations with PyECLOUD code for the dipole, quadrupole, sextupole magnets, and the warm (drift) section of the EIC hadron storage ring. PyECLOUD is an EC simulation code developed by CERN. The code has been validated and used to study EC in the LHC, SPS, and PS. To eliminate EC buildup, the sources of electrons must be minimized. First, we will reduce the production of electrons due to residual gas ionization by specifying the maximum gas density, which requires a vacuum chamber with low electron-stimulated desorption (ESD) yields and a sufficient (preferably distributed) pumping speed. Second, we should reduce the secondary electrons with a lower secondary electron yield (SEY) material.

43 PARTICLE ACCELERATORS↗
NASA NTRS1999

Flight Testing of a Cryogenic Capillary Pumped Loop

This paper describes the flight test results of the fifth generation cryogenic capillary pumped loop (CCPL-5) which flew on the Space Shuttle STS-95 in October of 1998 as part of the CRYOTSU Flight Experiment. This flight was the first in-space demonstration of the CCPL, a lightweight heat transport and thermal switching device for future integrated cryogenic bus systems. The CCPL-5 utilized nitrogen as the working fluid and operated between 80K and 110K. Flight results indicated excellent performance of the CCPL-5 under zero-G environment. The CCPL could start from a supercritical condition in all tests, and the loop operating temperature could be tightly controlled regardless of changes in the heat load and/or the sink temperature. In addition, the loop demonstrated successful operation with a heat load of 0.5 W as well as with parasitic heat loads alone.

Ku, Jentung↗
NASA NTRS1984

Long term cryogen storage in space - A feasibility study

The conduction of future space missions will require the long term storage of cryogens in space. Such missions will be performed in connection with the employment of the long life version of the Shuttle Infrared Telescope Facility (SIRTF), the Large Deployable Reflector (LDR), Gravity Probe B (GPB), Orbital Transfer Vehicles (OTVs), and the Space Station. The feasibility of building a long lifetime helium Dewar for space applications has been studied. A feasible concept was to be developed for a Shuttle-launchable 3-yr lifetime Dewar. The developed concept was to be used for the identification of key technology development areas. The desired long lifetime requires the minimization of the parasitic heat loads. It has been found that the majority of the parasitic heat load can be provided by the fiberglass-epoxy tension band supports. A new type of support system, the Passive Orbital Disconnect Strut (PODS), was developed to reduce the considered contribution to the heat load.

Kittel, P.↗