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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

Megawatt Class Nuclear Space Power Systems (MCNSPS) conceptual design and evaluation report. Volume 4: Concepts selection, conceptual designs, recommendations

A study was conducted by NASA Lewis Research Center for the Triagency SP-100 program office. The objective was to determine which reactor, conversion and radiator technologies would best fulfill future Megawatt Class Nuclear Space Power System Requirements. The requirement was 10 megawatts for 5 years of full power operation and 10 years system life on orbit. A variety of liquid metal and gas cooled reactors, static and dynamic conversion systems, and passive and dynamic radiators were considered. Four concepts were selected for more detailed study: (1) a gas cooled reactor with closed cycle Brayton turbine-alternator conversion with heatpipe and pumped tube fin rejection, (2) a Lithium cooled reactor with a free piston Stirling engine-linear alternator and a pumped tube-fin radiator,(3) a Lithium cooled reactor with a Potassium Rankine turbine-alternator and heat pipe radiator, and (4) a Lithium cooled incore thermionic static conversion reactor with a heat pipe radiator. The systems recommended for further development to meet a 10 megawatt long life requirement are the Lithium cooled reactor with the K-Rankine conversion and heat pipe radiator, and the Lithium cooled incore thermionic reactor with heat pipe radiator.

Wetch, J. R.↗
NASA NTRS1970

Radiant cooling

Passive radiant cooler design involving patch cone radiative coupling calculation based on specular image technique

Annable, R. V.↗
NASA NTRS1988

Megawatt Class Nuclear Space Power Systems (MCNSPS) conceptual design and evaluation report. Volume 1: Objectives, summary results and introduction

The objective was to determine which reactor, conversion, and radiator technologies would best fulfill future Megawatt Class Nuclear Space Power System Requirements. Specifically, the requirement was 10 megawatts for 5 years of full power operation and 10 years systems life on orbit. A variety of liquid metal and gas cooled reactors, static and dynamic conversion systems, and passive and dynamic radiators were considered. Four concepts were selected for more detailed study. The concepts are: a gas cooled reactor with closed cycle Brayton turbine-alternator conversion with heat pipe and pumped tube-fin heat rejection; a lithium cooled reactor with a free piston Stirling engine-linear alternator and a pumped tube-fin radiator; a lithium cooled reactor with potassium Rankine turbine-alternator and heat pipe radiator; and a lithium cooled incore thermionic static conversion reactor with a heat pipe radiator. The systems recommended for further development to meet a 10 megawatt long life requirement are the lithium cooled reactor with the K-Rankine conversion and heat pipe radiator, and the lithium cooled incore thermionic reactor with heat pipe radiator.

Wetch, J. R.↗
NASA NTRS1991

The transient gamma-ray spectrometer

The authors describe the Transient Gamma-Ray Spectrometer (TGRS) to be flown onboard the WIND spacecraft. This instrument is designed to detect cosmic gamma-ray bursts over the energy range of 20 keV to 10 MeV with an expected spectroscopic resolution of 2 keV at 1 MeV (E/Delta-E = 500). The active detection element is a 215-cu cm high-purity n-type Ge crystal cooled to cryogenic temperatures by a passive radiative cooler. The geometric field of view (FOV) defined by the cooler is 170 deg FWFM. Burst data are stored directly in an onboard 2.75-Mb burst memory with an absolute timing accuracy of +/-1.5 ms. This capacity is sufficient to store the entire spectral data set of all but the largest bursts. In addition to burst measurements, the instrument will also study solar flares, search for possible diffuse background lines, and monitor the 511-keV positron annihilation radiation from the galactic center. The experiment is scheduled to be launched on a Delta II launch vehicle from Cape Canaveral on December 31, 1992.

Owens, A.↗
NASA NTRS2019

Radiative Heating Rates Computed With Clouds Derived From Satellite‐Based Passive and Active Sensors and their Effects on Generation of Available Potential Energy

Radiative heating rates computed with cloud properties derived from passive and active sensors are investigated. Zonal monthly radiative heating rate anomalies computed using both active and passive sensors show that larger variability in longwave cooling exists near the tropical tropopause and near the top of the boundary layer between ~50°N to ~50°S. Aerosol variability contributes to increases in shortwave heating rate variability. When zonal monthly mean cloud effects on the radiative heating rate computed with both active and passive sensors and those computed with passive sensor only are compared, the latter shows cooling and heating peaks corresponding to cloud top and base height ranges used for separating cloud types. The difference of these two sets of cloud radiative effect on heating rates in the middle to upper troposphere is larger than the radiative heating rate uncertainty estimated based on the difference of two active sensor radiative heating rate profile data products. In addition, radiative heating rate contribution to generation of eddy available potential energy is also investigated. Although radiation contribution to generation of eddy available potential energy averaged over a year and the entire globe is small, radiation increases the eddy available potential energy in the northern hemisphere during summer. Two key elements that longwave radiation contribute to the generation of eddy potential energy are (1) longitudinal temperature gradient in the atmosphere associated with land and ocean surface temperatures contrasts and absorption of longwave radiation emitted by the surface and (2) cooling near the cloud top of stratocumulus clouds.

Kato, Seiji↗
NASA NTRS2009

Thermal Design and Analysis of a Multi-Stage 30K Radiative Cooling System for EPIC

The Experimental Probe of Inflationary Cosmology (EPIC) is an implementation of the NASA Einstein Inflation Probe mission, to answer questions about the physics of Inflation in the early Universe by measuring the polarization of the Cosmic Microwave Background (CMB). The mission relies on a passive cooling system to cool the enclosure of a telescope to 30 K; a cryocooler then cools this enclosure to 18 K and the telescope to 4 K. Subsequently, an adiabatic demagnetization refrigerator further cools a large focal plane to approx.100 mK. For this mission, the telescope has an aperture of 1.4 m, and the spacecraft's symmetry axis is oriented approx. 45 degrees relative to the direction of the sun. The spacecraft will be spun at approx. 0.5 rpm around this axis, which then precesses on the sky at 1 rph. The passive system must both supply the necessary cooling power for the cryocooler and meet demanding temperature stability requirements. We describe the thermal design of a passive cooling system consisting of four V-groove radiators for shielding of solar radiation and cooling the telescope to 30 K. The design realizes loads of 20 and 68 mW at the 4 K and 18 K stages on the cooler, respectively. A lower cost option for reaching 40 K with three V-groove radiators is also described. The analysis includes radiation coupling between stages of the radiators and sunshields, and parasitic conduction in the bipod support, harnesses, and ADR leads. Dynamic effects are also estimated, including the very small variations in temperature due to the scan motion of the spacecraft.

Space↗
NASA NTRS1995

Mechanical Pumped Cooling Loop for Spacecraft Thermal Control

Passive elements such as surface finishes, insulation, thermostats, radiators and louvers help keep spacecraft temperatures within the required limits. Mechanically pumped, single-phase cooling loops have been used for short-term applications such as the Space Shuttle. They are flexible, robust, and allow thermal autonomy. A lack of design heritage limits their use to shorter missions .

passive↗
NASA NTRS1974

Very high resolution radiometer

A primary sensor used in environmental and earth-resource observation, the Very High Resolution Radiometer (VHRR) was designed for use on the ITOS D series spacecraft. The VHRR provides a 0.47 mile resolution made possible with a mercury-cadmium-telluride detector cooled to approximately 105 K by a passive radiator cooler. The components of this system are described. The optical subsystem of the VHRR consists of a scanning mirror, a Dall-Kirkham telescope, a dichroic beam splitter, relay lenses, spectral filters, and an IR detector. Signal electronics amplify and condition the signals from the infrared and visible light detector. Sync generator electronics provides the necessary time signals. Scan-drive electronics is used for commutation of the motor winding, velocity, and phase control. A table lists the performance parameters of the VHRR.

Aronson, A. I.↗
NASA NTRS2018

The Origins Space Telescope Cryo-Thermal Architecture

The Origins Space Telescope (OST) is studied as a future Mid- and Far-Infrared flagship-class observatory. OST will cover the wavelength range from 6 to 600 microns. To reach the sky background for 200-micron wavelengths temperatures of 4 degrees Kelvin or lower are required. To achieve this low temperature active cooling is required, along with passive shielding and passive radiation to deep space. Currently two concepts are being studied: Concept 1 with a 9-meter-diameter primary and a suite of 5 extremely capable instruments providing both imaging and spectroscopy over the entire wavelength range. Concept 2 is a more modest sized telescope with a collecting area equivalent to a 5-meter primary, fewer deployments and 3 or 4 instruments also covering the entire wavelength range for imaging and spectroscopy, although with somewhat reduced spectroscopic resolution, and somewhat slower mapping speed. This paper will describe OST Concept 2's cryogenic thermal architecture and thermal model results.

Space Telescope↗
NASA NTRS2016

James Webb Space Telescope Core 2 Test - Cryogenic Thermal Balance Test of the Observatorys Core Area Thermal Control Hardware

The James Webb Space Telescope (JWST), successor to the Hubble Space Telescope, will be the largest astronomical telescope ever sent into space. To observe the very first light of the early universe, JWST requires a large deployed 6.5-meter primary mirror cryogenically cooled to less than 50 Kelvin. Three scientific instruments are further cooled via a large radiator system to less than 40 Kelvin. A fourth scientific instrument is cooled to less than 7 Kelvin using a combination pulse-tube Joule-Thomson mechanical cooler. Passive cryogenic cooling enables the large scale of the telescope which must be highly folded for launch on an Ariane 5 launch vehicle and deployed once on orbit during its journey to the second Earth-Sun Lagrange point. Passive cooling of the observatory is enabled by the deployment of a large tennis court sized five layer Sunshield combined with the use of a network of high efficiency radiators. A high purity aluminum heat strap system connects the three instrument's detector systems to the radiator systems to dissipate less than a single watt of parasitic and instrument dissipated heat. JWST's large scale features, while enabling passive cooling, also prevent the typical flight configuration fully-deployed thermal balance test that is the keystone of most space missions' thermal verification plans. This paper describes the JWST Core 2 Test, which is a cryogenic thermal balance test of a full size, high fidelity engineering model of the Observatory's 'Core' area thermal control hardware. The 'Core' area is the key mechanical and cryogenic interface area between all Observatory elements. The 'Core' area thermal control hardware allows for temperature transition of 300K to approximately 50 K by attenuating heat from the room temperature IEC (instrument electronics) and the Spacecraft Bus. Since the flight hardware is not available for test, the Core 2 test uses high fidelity and flight-like reproductions.

JWST Thermal Core 2 Test↗
NASA NTRS2008

Optical Coating Performance and Thermal Structure Design for Heat Reflectors of JWST Electronic Control Unit

The James Webb Space Telescope (JWST) consists of an infrared-optimized Optical Telescope Element (OTE) that is cooled down to 40 degrees Kelvin. A second adjacent component to the OTE is the Integrated Science Instrument Module, or ISIM. This module includes the electronic compartment, which provides the mounting surfaces and ambient thermally controlled environment for the instrument control electronics. Dissipating the 200 watts generated from the ISIM structure away from the OTE is of paramount importance so that the spacecraft's own heat does not interfere with the infrared light detected from distant cosmic sources. This technical challenge is overcome by a thermal subsystem unit that provides passive cooling to the ISIM control electronics. The proposed design of this thermal radiator consists of a lightweight structure made out of composite materials and low-emittance metal coatings. In this paper, we will present characterizations of the coating emittance, bidirectional reflectance, and mechanical structure design that will affect the performance of this passive cooling system.

Quijada, Manuel A.↗
NASA NTRS1993

Thermal systems design and analysis for a 10 K Sorption Cryocooler flight experiment

The design, analysis and predicted performance of the Brilliant Eyes Ten-Kelvin Sorption Cryocooler Experiment (BETSCE) is described from a thermal perspective. BETSCE is a shuttle side-wall mounted cryogenic technology demonstration experiment planned for launch in November 1994. BETSCE uses a significant amount of power (about 500 W peak) and the resultant heat must be rejected passively with radiators, as BETSCE has no access to the active cooling capability of the shuttle. It was a major challenge to design and configure the individual hardware assemblies, with their relatively large radiators, to enable them to reject their heat while satisfying numerous severe shuttle-imposed constraints. This paper is a useful case study of a small shuttle payload that needs to reject relatively high heat loads passively in a highly constrained thermal environment. The design approach described is consistent with today's era of 'faster, better, cheaper' small-scale space missions.

Bhandari, Pradeep↗
NASA NTRS2013

Thermal Infrared Sensor (TIRS) Instrument Thermal Subsystem Design and Lessons Learned

The Thermal Infrared Sensor (TIRS) is one of two instruments on the Landsat Data Continuity Mission (LDCM), which is scheduled to launch in February of 2013. The TIRS instrument was officially added to the mission later in the flow, which led to a highly aggressive schedule that became one of the main drivers during instrument development. The thermal subsystem design of the TIRS Sensor Unit is comprised of five thermal zones which range in temperature from less than 43 Kelvin to 330 Kelvin. Most zones are proportional heater controlled, and all are within a volume of 35 cu.ft. A two-stage cryocooler is used to cool the "cold stage" including three QWIP detectors to less than 43 Kelvin, and cool the "warm stage" to 105 Kelvin. The excess power dissipation from the cryocooler is rejected via ammonia transport heat pipes to a dedicated Cryocooler Radiator with embedded ammonia heat pipes. The cryogenic subsystem includes a series of shells used to radiatively and conductively isolate the cold stage from the warmer surroundings. The Optical System (telescope) is passively cooled to 180-190 Kelvin using a "thermal link" (comprised of a Flexible Conductive Thermal Strap and an APG Bar) which couples the telescope stage to a dedicated radiator with embedded ethane heat pipes. The Scene Select Mechanism, which is responsible for moving the Scene Select Mirror to three distinct positions (including Nadir, Space, and On-board Black Body Calibrator pointing), runs nominally at 278 Kelvin and is thermally isolated from the cryogenic thermal zones. The On-board Black Body Calibrator requires a dedicated radiator which allows for a temperature range of 260-330 Kelvin at the Source. The detectors are powered by the FPE Box, which is mounted to the nadir external surface of the composite honeycomb structure. There are two additional electronics boxes which are wet-mounted directly to the spacecraft shear panel, the Main Electronics Box and Cryocooler Electronics Box; thermal control of these boxes is the responsibility of Orbital Sciences Corporation, the spacecraft developer. The TIRS thermal subsystem design was successfully verified during months of testing campaign, from component & subsystem level to two instrument-level thermal vacuum tests. The Instrument, despite an aggressive schedule, was delivered to the spacecraft vendor in February of 2012 and is currently undergoing the final stages of spacecraft environmental testing in preparation for launch.

Otero, Veronica↗
NASA NTRS2018

Multi-Angle Imager for Aerosols Thermal Control System

The Multi-Angle Imager for Aerosols (MAIA) Thermal Control System is a NASA funded instrument that will collect data to help characterize airborne particulate matter over a number of population centers across the globe using multi-angle spectropolarimetric imagery. The data collected by MAIA will facilitate assessments of the impacts of different types of particulate matter on adverse health outcomes. MAIA is a hosted payload meant to operate in a near-circular sun-synchronous polar orbit, with a mean altitude between 600 km and 850 km. The nominal on-orbit mission design lifetime is three years. Temperature control of the MAIA instrument is accomplished with a combination of passive radiators and heaters. The focal plane module (FPM) is cooled to ≤ 235K with a disc shaped radiator that faces the anti-sun side of the sun-synchronous orbit. The temperature of the MAIA cameras and associated electronics is controlled with a cylindrical shaped radiator that projects a near constant area in the nadir direction as the cameras rotate. A noteworthy feature of the MAIA thermal control system design is the novel, low cost, rotationally articulating thermal strap used to transfer heat from the FPMs to their associated FPM Radiator. The strap spans one of the axes of rotation, sweeping out an arc of approximately 60° as the instrument operates. A prototype of the articulating thermal strap was life tested to 260,000 cycles with no signs of significant degradation. An overview of the MAIA thermal control system baseline design is presented, with focus on its novel aspects, including life testing of the prototype articulating thermal strap. In addition, a discussion of the considerations involved in designing a thermal control system for a hosted instrument is provided.

Rosas, Rogelio↗
NASA NTRS2026

Assessing the Adhesion of Nanofibrous PVDF-HFP as Passive Thermal Control Coatings for the Extraterrestrial Storage of Cryogenic Propellants

Cryogenic propellant passive thermal management techniques in space are essential complements to active thermal management methods, helping reduce absolute dependence on the latter. Particularly, materials with superior optical properties (high solar reflectance and infrared emittance) would be beneficial for passive thermal control to reject most of the incident solar radiation and enhance thermal emission from the storage tank, leading to self-cooling. We developed a nanofibrous, porous passive thermal control material by electrospinning polyvinylidene fluoride-co-hexafluoropropylene (PVDF-HFP) co-polymer onto an aluminum foil substrate. The material exhibits a very high solar reflectance (>99%) and infrared emittance (~86%) when characterized in the ultraviolet-visible-near infrared and mid-infrared wavelengths. This study focuses on the adhesion of the nanofibrous PVDF-HFP to the substrate. We incorporated in-lab fabricated and commercial adhesives, under normal room conditions, between the nanofibers and the aluminum foil substrate to improve their adhesion. Some samples were also exposed to space-like conditions – vacuum and thermal cycling, to investigate their effects on the strengths of adhesives. Subsequently, adhesion peel tests were conducted on all the samples to assess the peel strengths of the bonds. Peel test strengths after space environment conditioning of the adhesives were generally higher than those observed in the pristine, unaged samples. The results are discussed in this study.

Chieloka Ibekwe↗
NASA NTRS2016

James Webb Space Telescope Integrated Science Instrument Module Thermal Vacuum Thermal Balance Test Campaign at NASA's Goddard Space Flight Center

The James Webb Space Telescope is a large infrared telescope with a 6.5-meter primary mirror, designed as a successor to the Hubble Space Telescope when launched in 2018. Three of the four science instruments contained within the Integrated Science Instrument Module (ISIM) are passively cooled to their operational temperature range of 36K to 40K with radiators, and the fourth instrument is actively cooled to its operational temperature of approximately 6K. Thermal-vacuum testing of the flight science instruments at the ISIM element level has taken place in three separate highly challenging and extremely complex thermal tests within a gaseous helium-cooled shroud inside Goddard Space Flight Centers Space Environment Simulator. Special data acquisition software was developed for these tests to monitor over 1700 flight and test sensor measurements, track over 50 gradients, component rates, and temperature limits in real time against defined constraints and limitations, and guide the complex transition from ambient to final cryogenic temperatures and back. This extremely flexible system has proven highly successful in safeguarding the nearly $2B science payload during the 3.5-month-long thermal tests. Heat flow measurement instrumentation, or Q-meters, were also specially developed for these tests. These devices provide thermal boundaries o the flight hardware while measuring instrument heat loads up to 600 mW with an estimated uncertainty of 2 mW in test, enabling accurate thermal model correlation, hardware design validation, and workmanship verification. The high accuracy heat load measurements provided first evidence of a potentially serious hardware design issue that was subsequently corrected. This paper provides an overview of the ISIM-level thermal-vacuum tests and thermal objectives; explains the thermal test configuration and thermal balances; describes special measurement instrumentation and monitoring and control software; presents key test thermal results; lists problems encountered during testing and lessons learned.

JWST ISIM Thermal↗
NASA NTRS1972

Completely modular Thermionic Reactor Ion Propulsion System (TRIPS)

The nuclear reactor powered ion propulsion system described is an advanced completely modularized system which lends itself to development of prototype and/or flight type components without the need for complete system tests until late in the development program. This modularity is achieved in all of the subsystems and components of the electric propulsion system including (1) the thermionic fuel elements, (2) the heat rejection subsystem (heat pipes), (3) the power conditioning modules, and (4) the ion thrusters. Both flashlight and external fuel type in-core thermionic reactors are considered as the power source. The thermionic fuel elements would be useful over a range of reactor power levels. Electrical heated acceptance testing in their flight configuration is possible for the external fuel case. Nuclear heated testing by sampling methods could be used for acceptance testing of flashlight fuel elements. The use of heat pipes for cooling the collectors and as a means of heat transport to the radiator allows early prototype or flight configuration testing of a small module of the heat rejection subsystem as opposed to full scale liquid metal pumps and radiators in a large vacuum chamber. The power conditioner (p/c) is arranged in modules with passive cooling.

Peelgren, M. L.↗
NASA NTRS1972

Completely modular thermionic reactor ion propulsion system /trips/.

Description of a nuclear reactor powered ion propulsion system which is an advanced completely modularized system lending itself to development of prototype and/or flight type components without the need for complete system tests until late in the development program. This modularity is achieved in all of the subsystems and components of the electric propulsion system, including the thermionic fuel elements, the heat rejection subsystem (head pipes), the power conditioning modules, and the ion thrusters. The use of heat pipes for cooling the collectors and as a means of heat transport to the radiator allows early prototype or flight configuration testing of a small module of the heat rejection subsystem as opposed to full scale liquid metal pumps and radiators in a large vacuum chamber. The power conditioner (p/c) is arranged in modules with passive cooling which allows complete prototype testing. The ion engines are typically matched with one or more p/c modules and are the same size for any power level propulsion system of interest.

Peelgren, M. L.↗