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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Pump and Flow Control Subassembly of Thermal Control Subsystem for Photovoltaic Power Module

The pump and flow control subassembly (PFCS) is an orbital replacement unit (ORU) on the Space Station Freedom photovoltaic power module (PVM). The PFCS pumps liquid ammonia at a constant rate of approximately 1170 kg/hr while providing temperature control by flow regulation between the radiator and the bypass loop. Also, housed within the ORU is an accumulator to compensate for fluid volumetric changes as well as the electronics and firmware for monitoring and control of the photovoltaic thermal control system (PVTCS). Major electronic functions include signal conditioning, data interfacing and motor control. This paper will provide a description of each major component within the PFCS along with performance test data. In addition, this paper will discuss the flow control algorithm and describe how the nickel hydrogen batteries and associated power electronics will be thermally controlled through regulation of coolant flow to the radiator.

Motil, Brian↗
NASA NTRS1978

Orbiter thermal control subsystem development

Selected major hardware components and subsystems of the shuttle orbiter are described. These include the hardware configurations, development problems and solutions, performance requirements and attainment, major hardware fabrication problems and solutions, development test objectives and implementation, and utilization of test results for hardware refinements. The subsystems selected for discussion are the space radiator, the ammonia boiler, the flash evaporator, and avionics cold plates.

Laubach, G. E.↗
NASA NTRS1992

Space Station Furnace Facility. Volume 2: Summary of technical reports

The Space Station Furnace Facility (SSFF) is a modular facility for materials research in the microgravity environment of the Space Station Freedom (SSF). The SSFF is designed for crystal growth and solidification research in the fields of electronic and photonic materials, metals and alloys, and glasses and ceramics, and will allow for experimental determination of the role of gravitational forces in the solidification process. The facility will provide a capability for basic scientific research and will evaluate the commercial viability of low-gravity processing of selected technologically important materials. In order to accommodate the furnace modules with the resources required to operate, SSFF developed a design that meets the needs of the wide range of furnaces that are planned for the SSFF. The system design is divided into subsystems which provide the functions of interfacing to the SSF services, conditioning and control for furnace module use, providing the controlled services to the furnace modules, and interfacing to and acquiring data from the furnace modules. The subsystems, described in detail, are as follows: Power Conditioning and Distribution Subsystem; Data Management Subsystem; Software; Gas Distribution Subsystem; Thermal Control Subsystem; and Mechanical Structures Subsystem.

Source record↗
NASA NTRS1986

Independent Orbiter Assessment (IOA): Analysis of the electrical power distribution and control/electrical power generation subsystem

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Electrical Power Distribution and Control (EPD and C)/Electrical Power Generation (EPG) hardware. The EPD and C/EPG hardware is required for performing critical functions of cryogenic reactant storage, electrical power generation and product water distribution in the Orbiter. Specifically, the EPD and C/EPG hardware consists of the following components: Power Section Assembly (PSA); Reactant Control Subsystem (RCS); Thermal Control Subsystem (TCS); Water Removal Subsystem (WRS); and Power Reactant Storage and Distribution System (PRSDS). The IOA analysis process utilized available EPD and C/EPG hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Patton, Jeff A.↗
NASA NTRS1990

Electrical power system WP-04

Viewgraphs on Space Station Freedom Electrical Power System (EPS) WP-40 are presented. Topics covered include: key EPS technical requirements; photovoltaic power module systems; solar array assembly; blanket containment box and box positioning subassemblies; solar cell; bypass diode assembly; Kapton with atomic oxygen resistant coating; sequential shunt unit; gimbal assembly; energy storage subsystem; thermal control subsystem; direct current switching unit; integrated equipment assembly; PV cargo element; PMAD system; and PMC and AC architecture.

Nored, Donald L.↗
NASA NTRS1972

Advanced extravehicular protective systems study, volume 1

An appraisal was made of advanced portable and emergency life support systems concepts for space station, space shuttle, lunar base, and Mars EVA missions. Specifications are given, and the methodology is described. Subsystem studies and systems integration efforts are summarized. Among the conclusions are the following: (1) For long duration missions, a configuration incorporating a regenerable CO2 control subsystem and a thermal control subsystem utilizing a minimum of expendables decreases the vehicle penalty of present configurations. (2) For shorter duration missions, a configuration incorporating an expendable water thermal control subsystem is the most competitive subsystem; regenerable CO2 control subsystems if properly developed are competitive with nonregenerable counterparts. (3) The CO2 reduction and oxygen reclamation withing the parent vehicle is only competitive when there are three or more parent vehicle resupply periods. (4) For long duration emergency systems of one hour or more, inherent redundancy within the primary configuration to provide emergency thermal control is the most competitive approach.

Sutton, J. G.↗
NASA NTRS1986

Independent Orbiter Assessment (IOA): Analysis of the electrical power generation/fuel cell powerplant subsystem

Results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Electrical Power Generation (EPG)/Fuel Cell Powerplant (FCP) hardware. The EPG/FCP hardware is required for performing functions of electrical power generation and product water distribution in the Orbiter. Specifically, the EPG/FCP hardware consists of the following divisions: (1) Power Section Assembly (PSA); (2) Reactant Control Subsystem (RCS); (3) Thermal Control Subsystem (TCS); and (4) Water Removal Subsystem (WRS). The IOA analysis process utilized available EPG/FCP hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Brown, K. L.↗
NASA NTRS2011

Packing the PLSS

NASA Engineers design spacesuits for ultimate protection and functionality in the extreme environment of space. The spacesuit is often referred to as a "personal spacecraft" because it provides the astronaut with everything he or she needs to survive and work in space outside of the vehicle or habitat. The systems within the spacesuit include the pressure garment system (PGS), the Portable Life Support System (PLSS), and the power, avionics, and software (PAS) system. These elements are necessary to protect crewmembers and allow them to work effectively in the pressure and temperature extremes of space environments. Development of the spacesuit system is necessary to support future human extravehicular exploration activities to Lunar, Martian, microgravity, and possibly other space destinations. Although all the systems that makeup the space suit are important, the PLSS is one of the most complex. The PLSS provides the life support needed by the astronaut and consists of the oxygen (O2) subsystem, ventilation subsystem, and thermal control subsystem. Within each subsystem, there are many different components, a few of which are explained as follows. The oxygen tanks hold the oxygen that the crewmember uses to breath and pressurizes the suit. The primary oxygen tank is responsible during normal operations and the secondary oxygen tank kicks on in the case of an emergency. The Rapid Cycle Amine (RCA) canister is used to remove the carbon dioxide (CO2) and extra humidity in the crewmember's ventilation/breathing gas. The fan moves the oxygen around the suit. Suit Water Membrane Evaporator (SWME) is used within the thermal control loop to cool the water that is used to maintain a comfortable temperature for both the crew member and the other equipment inside the suit. Another component is the battery, which supplies the power needed to operate all these and the many other pieces. The battery is one of the biggest and heavies components within the PLSS. These are just a few of the components that encompass the PLSS. Each component has a weight and a certain volume that the NASA Engineers must take into account when building the PLSS, because the weight and volumes affect the crewmembers center of gravity (CG). [See the Notes Section for the link to an Apollo video that demonstrates the issues some of the crewmembers had picking up tools and dealing with center of gravity/tools on the surface of the Moon.] In this activity, students will simulate engineering design techniques that NASA Engineers and Designers are currently implementing to configuring the components within the PLSS. Through testing, students will consider the comfort, mobility, and center of gravity for their test subjects and how that changes after adjusting the placement of their simulated PLSS components.

Jennings, Mallory↗
NASA NTRS1979

A thermal control approach for a solar electric propulsion thrust subsystem

A thrust subsystem thermal control design is defined for a Solar Electric Propulsion System (SEPS) proposed for the comet Halley Flyby/comet Tempel 2 rendezvous mission. A 114 node analytic model, developed and coded on the systems improved numerical differencing analyzer program, was employed. A description of the resulting thrust subsystem thermal design is presented as well as a description of the analytic model and comparisons of the predicted temperature profiles for various SEPS thermal configurations that were generated using this model. It was concluded that: (1) a BIMOD engine system thermal design can be autonomous; (2) an independent thrust subsystem thermal design is feasible; (3) the interface module electronics temperatures can be controlled by a passive radiator and supplementary heaters; (4) maintaining heat pipes above the freezing point would require an additional 322 watts of supplementary heating power for the situation where no thrusters are operating; (5) insulation is required around the power processors, and between the interface module and the avionics module, as well as in those areas which may be subjected to solar heating; and (6) insulation behind the heat pipe radiators is not necessary.

Maloy, J. E.↗
NASA NTRS1974

Applications Technology Satellite ATS-6 in orbit checkout report

The activities of the ATS-6 spacecraft for the checkout period of approximately four weeks beginning May 30, 1974 are described, along with the results of a performance evaluation of its subsystems and components. The following specific items are discussed: (1) subsystem requirements/specifications and in-orbit performance summary; (2) flight chronology; (3) spacecraft description; (4) structural/deployment subsystems; (5) electrical power subsystem; (6) thermal control subsystem; (7) telemetry and command subsystems; (8) attitude control subsystem; (9) spacecraft propulsion subsystem; (10) communication subsystem; and (12) experiment subsystem.

Moore, W.↗
NASA NTRS1974

Earth Observatory Satellite system definition study. Report 5: System design and specifications. Volume 3: General purpose spacecraft segment and module specifications

The specifications for the Earth Observatory Satellite (EOS) general purpose aircraft segment are presented. The satellite is designed to provide attitude stabilization, electrical power, and a communications data handling subsystem which can support various mission peculiar subsystems. The various specifications considered include the following: (1) structures subsystem, (2) thermal control subsystem, (3) communications and data handling subsystem module, (4) attitude control subsystem module, (5) power subsystem module, and (6) electrical integration subsystem.

Source record↗
NASA NTRS1976

Space telescope optical telescope assembly/scientific instruments. Phase B: -Preliminary design and program definition study; Volume 2A: Planetary camera report

Development of the F/48, F/96 Planetary Camera for the Large Space Telescope is discussed. Instrument characteristics, optical design, and CCD camera submodule thermal design are considered along with structural subsystem and thermal control subsystem. Weight, electrical subsystem, and support equipment requirements are also included.

Source record↗
NASA NTRS2001

Development of the Variable Emittance Thermal Suite for the Space Technology 5 Microsatellite

The advent of very small satellites, such as nano and microsatellites, logically leads to a requirement for smaller thermal control subsystems. In addition, the thermal control needs of the smaller spacecraft/instrument may well be different from more traditional situations. For example, power for traditional heaters may be very limited or unavailable, mass allocations may be severely limited, and fleets of nano/microsatellites will require a generic thermal design as the cost of unique designs will be prohibitive. Some applications may require significantly increased power levels while others may require extremely low heat loss for extended periods. Small spacecraft will have low thermal capacitance thus subjecting them to large temperature swings when either the heat generation rate changes or the thermal sink temperature changes. This situation, combined with the need for tighter temperature control, will present a challenging situation during transient operation. The use of "off-the-shelf" commercial spacecraft buses for science instruments will also present challenges. Older thermal technology, such as heaters, thermostats, and heat pipes, will almost certainly not be sufficient to meet the requirements of these new spacecraft/instruments. They are generally too heavy, not scalable to very small sizes, and may consume inordinate amounts of power. Hence there is a strong driver to develop new technology to meet these emerging needs. Variable emittance coatings offer an exciting alternative to traditional control methodologies and are one of the technologies that will be flown on Space Technology 5, a mission of three microsatellites designed to validate "enabling" technologies. Several studies have identified variable emittance coatings as applicable to a wide range of spacecraft, and to potentially offer substantial savings in mass and/or power over traditional approaches. This paper discusses the development of the variable emittance thermal suite for ST-5. More specifically, it provides a description of and the infusion and validation plans for the variable emittance coatings.

Douglas, Donya M.↗
NASA NTRS2014

Embedded Thermal Control for Spacecraft Subsystems Miniaturization

Optimization of spacecraft size, weight and power (SWaP) resources is an explicit technical priority at Goddard Space Flight Center. Embedded Thermal Control Subsystems are a promising technology with many cross cutting NSAA, DoD and commercial applications: 1.) CubeSatSmallSat spacecraft architecture, 2.) high performance computing, 3.) On-board spacecraft electronics, 4.) Power electronics and RF arrays. The Embedded Thermal Control Subsystem technology development efforts focus on component, board and enclosure level devices that will ultimately include intelligent capabilities. The presentation will discuss electric, capillary and hybrid based hardware research and development efforts at Goddard Space Flight Center. The Embedded Thermal Control Subsystem development program consists of interrelated sub-initiatives, e.g., chip component level thermal control devices, self-sensing thermal management, advanced manufactured structures. This presentation includes technical status and progress on each of these investigations. Future sub-initiatives, technical milestones and program goals will be presented.

Thermal Control↗
NASA NTRS1983

Space station needs, attributes, and architectural options study. Volume 2: Program options, architecture, and technology

Mission scenarios and space station architectures are discussed. Electrical power subsystems (EPS), environmental control and life support, subsystems (ECLSS), and reaction control subsystem (RCS) architectures are addressed. Thermal control subsystems, (TCS), guidance/navigation and control (GN and C), information management systems IMS), communications and tracking (C and T), and propellant transfer and storage systems architectures are discussed.

Source record↗
NASA NTRSDate not supplied

ISRU Pilot Excavator (IPEx) Technology Readiness Level 5 Design Overview

This paper details the mechanical and mechatronic design of the Technology Readiness Level (TRL)-5 In-Situ Resource Utilization (ISRU) Pilot Excavator (IPEx). IPEx is a robotic excavator designed for a technology demonstration of regolith mining in the lunar south -pole region. The novel design uses pairs of counter-acting excavation tools called bucket drums, that dig at the same time in opposing directions to reduce the reaction force needed, thereby enabling mining with a small, low-mass, robotic system. IPEx builds on the prior work of the Regolith Advanced Surface Systems Operations Robot (RASSOR), which is the TRL-4 implementation of this concept. The TRL-5 IPEx subsystems that are discussed in this paper include: Regolith Delivery Subsystem (RDS), Mobility Subsystem (MS), Cameras and Dust Mitigation Subsystem (CDMS), and Thermal Control Subsystem (TCS). Each subsystem is described in detail with rationale for design selections. Dust tolerance is a key feature for IPEx and this paper details a thermal control system with an actuated radiator cover and phase change material as well as camera modules with removable electrodynamic dust shields (EDS). Additional components such as actuators, wheels, and bucket drums are discussed in detail. Due to their complexity, the avionics and software subsystems will be discussed in a separate publication.

RASSOR↗
NASA NTRS2014

CFD Analysis of Thermal Control System Using NX Thermal and Flow

The Thermal Control Subsystem (TCS) is a key part of the Advanced Plant Habitat (APH) for the International Space Station (ISS). The purpose of this subsystem is to provide thermal control, mainly cooling, to the other APH subsystems. One of these subsystems, the Environmental Control Subsystem (ECS), controls the temperature and humidity of the growth chamber (GC) air to optimize the growth of plants in the habitat. The TCS provides thermal control to the ECS with three cold plates, which use Thermoelectric Coolers (TECs) to heat or cool water as needed to control the air temperature in the ECS system. In order to optimize the TCS design, pressure drop and heat transfer analyses were needed. The analysis for this system was performed in Siemens NX Thermal/Flow software (Version 8.5). NX Thermal/Flow has the ability to perform 1D or 3D flow solutions. The 1D flow solver can be used to represent simple geometries, such as pipes and tubes. The 1D flow method also has the ability to simulate either fluid only or fluid and wall regions. The 3D flow solver is similar to other Computational Fluid Dynamic (CFD) software. TCS performance was analyzed using both the 1D and 3D solvers. Each method produced different results, which will be evaluated and discussed.

CFD↗
NASA NTRS2011

Impacts of an Ammonia Leak on the Cabin Atmosphere of the International Space Station

Toxic chemical release into the cabin atmosphere is one of the three major emergency scenarios identified on the International Space Station (ISS). The release of anhydrous ammonia, the coolant used in the U.S. On-orbit Segment (USOS) External Active Thermal Control Subsystem (EATCS), into the ISS cabin atmosphere is one of the most serious toxic chemical release cases identified on board ISS. The USOS Thermal Control System (TCS) includes an Internal Thermal Control Subsystem (ITCS) water loop and an EATCS ammonia loop that transfer heat at the interface heat exchanger (IFHX). Failure modes exist that could cause a breach within the IFHX. This breach would result in high pressure ammonia from the EATCS flowing into the lower pressure ITCS water loop. As the pressure builds in the ITCS loop, it is likely that the gas trap, which has the lowest maximum design pressure within the ITCS, would burst and cause ammonia to enter the ISS atmosphere. It is crucial to first characterize the release of ammonia into the ISS atmosphere in order to develop methods to properly mitigate the environmental risk. This paper will document the methods used to characterize an ammonia leak into the ISS cabin atmosphere. A mathematical model of the leak was first developed in order to define the flow of ammonia into the ISS cabin atmosphere based on a series of IFHX rupture cases. Computational Fluid Dynamics (CFD) methods were then used to model the dispersion of the ammonia throughout the ISS cabin and determine localized effects and ventilation effects on the dispersion of ammonia. Lastly, the capabilities of the current on-orbit systems to remove ammonia were reviewed and scrubbing rates of the ISS systems were defined based on the ammonia release models. With this full characterization of the release of ammonia from the USOS TCS, an appropriate mitigation strategy that includes crew and system emergency response procedures, personal protection equipment use, and atmosphere monitoring and scrubbing hardware can be established.

Duchesne, Stephanie M.↗