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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 NTRS2019

James Webb Space Telescope Optical Telescope Element / Integrated Science Instrument Module (OTIS) Cryogenic Vacuum Test Part I: Thermal Architecture

The James Webb Space Telescope (JWST), set to launch in early 2019, is currently undergoing a series of system-level environmental tests to verify its workmanship and end-to-end functionality. As part of this series, the Optical Telescope Element and Integrated Science Instrument Module (OTIS) Cryo-Vacuum (CV) test, the most complex cryogenic test executed to date by NASA, was completed in 2017 at the Johnson Space Center's Chamber A facility. The OTIS CV test was intended as a comprehensive test of the integrated instrument and telescope systems to fully understand its optical, structural, and thermal performance within its intended flight environment. Due to its complexity, extensive pre-test planning was required to ensure payload safety and compliance with all limits and constraints. A comprehensive thermal analysis campaign was undertaken to predict the behavior of the test payload during its transition to cryogenic temperatures and back to ambient, and exhaustive preparations for off-nominal scenarios were studied due to the increased possibility of unanticipated events during the 100-day test length. This current four-part online course for the NASA Engineering and Safety Center (NESC) Academy gives a thorough overview of the mutiple years of preparation executed by hundreds of individuals to ensure the success of the JWST OTIS CV Test. Part I will introduce the major components of James Webb and specifically the OTIS CV Test thermal architecture. Part II will encompass the extensive thermal analysis performed to prepare for the test. Part III will talk about the preparations for off-nominal events: what analysis was done, and how steps were taken pre-test to anticipate unexpected circumstances and mitigate their impacts to hardware and test timeline. Part IV is a recap of the lessons learned from the thermal perspective for both the payload and ground support equipment (GSE) test conductors. It is hoped that the insight gained from the OTIS CV test campaign will benefit planning and execution for upcoming large cryogenic missions.

OTIS↗
NASA NTRS2019

James Webb Space Telescope Optical Telescope Element / Integrated Science Instrument Module (OTIS) Cryogenic Vacuum Test Part II: Thermal Analysis

The James Webb Space Telescope (JWST), set to launch in early 2019, is currently undergoing a series of system-level environmental tests to verify its workmanship and end-to-end functionality. As part of this series, the Optical Telescope Element and Integrated Science Instrument Module (OTIS) Cryo-Vacuum (CV) test, the most complex cryogenic test executed to date by NASA, was completed in 2017 at the Johnson Space Center's Chamber A facility. The OTIS CV test was intended as a comprehensive test of the integrated instrument and telescope systems to fully understand its optical, structural, and thermal performance within its intended flight environment. Due to its complexity, extensive pre-test planning was required to ensure payload safety and compliance with all limits and constraints. A comprehensive thermal analysis campaign was undertaken to predict the behavior of the test payload during its transition to cryogenic temperatures and back to ambient, and exhaustive preparations for off-nominal scenarios were studied due to the increased possibility of unanticipated events during the 100-day test length. This current four-part online course for the NASA Engineering and Safety Center (NESC) Academy gives a thorough overview of the multiple years of preparation executed by hundreds of individuals to ensure the success of the JWST OTIS CV Test. Part I will introduce the major components of James Webb and specifically the OTIS CV Test thermal architecture. Part II will encompass the extensive thermal analysis performed to prepare for the test. Part III will talk about the preparations for off-nominal events: what analysis was done, and how steps were taken pre-test to anticipate unexpected circumstances and mitigate their impacts to hardware and test timeline. Part IV is a recap of the lessons learned from the thermal perspective for both the payload and ground support equipment (GSE) test conductors. It is hoped that the insight gained from the OTIS CV test campaign will benefit planning and execution for upcoming large cryogenic missions.

Testing↗
NASA NTRS2018

A Hybrid Electrostatic Retarding Potential Analyzer for the Measurement of Plasmas at Extremely High Energy Resolution

Many space plasmas (especially electrons generated in planetary ionospheres) exhibit fine-detailed structures that are challenging to fully resolve with the energy resolution of typical space plasma analyzers (10% → 20%). While analyzers with higher resolution have flown, generally this comes at the expense of sensitivity and temporal resolution. We present a new technique for measuring plasmas with extremely high energy resolution through the combination of a top-hat Electrostatic Analyzer (ESA) followed by an internally mounted Retarding Potential Analyzer (RPA). When high resolutions are not required, the RPA is grounded, and the instrument may operate as a typical general-purpose plasma analyzer using its ESA alone. We also describe how such an instrument may use its RPA to remotely vary the geometric factor (sensitivity) of a top hat analyzer, as was performed on the New Horizons Solar Wind at Pluto and MAVEN SupraThermal and Thermal Ion Composition instruments. Finally, we present results from laboratory testing of our prototype, showing that this technique may be used to construct an instrument with 1.6% energy resolution, constant over all energies and angles.

Collinson, Glyn A.↗
NASA NTRS1986

Thermal infrared opacity of the Mars atmosphere

An analytical technique is presented for the deriving Mars atmosphere opacity from data furnished by the IR Thermal Mapper (IRTM) instruments on the Viking landers. Each IRTM features four telescopes for viewing Mars in six spectral bands and seven spatial fields of view. The present study focuses on 20 microns measurements of the thermal inertia, 0.3-3.0 microns data for the surface albedo, and, finally, the 7, 9, and 15 microns data for the actual opacity derivations. The model takes into account the atmospheric temperature profile, scattering by dust, the surface emissivity and the variation of the surface/atmosphere temperature discontinuity. Analyses are carried out of the possible error factors and the model is used to generate opacity histories for the two Viking Lander sites. The results, when compared with other astronomical data, indicate that the method is sufficiently accurate to use opacity data to analyze the processes of storm genesis and to study local storms.

Martin, T. Z.↗
NASA NTRS1991

A survey of spacecraft thermal design solutions

A review of activities at the NASA/Marshall Space Flight Center in the heat transfer and thermodynamics disciplines as well as attendant fluid mechanics, transport phenomena, and computer science applications is presented. Attention is focused on recent activities including the Hubble Space Telescope, and large space instruments, particularly telescope thermal control systems such as those flown aboard Spacelab 2 and the Astro missions. Emphasis is placed on defining the thermal control features, unique design schemes, and performance of selected programs. Results obtained both by ground testing and analytical means, as well as flight and postflight data are presented.

Humphries, R.↗
NASA NTRS1987

Shuttle payload bay thermal environments: Summary and conclusion report for STS Flights 1-5

The thermal data for the payload bay of the first five shuttle flights is summarized and the engineering evaluation of that data is presented. After a general discussion on mission profiles and vehicle configurations, the thermal design and flight instrumentation systems of the payload bay are described. The thermal flight data sources and a categorization of the data are then presented. A thermal flight data summarization section provides temperature data for the five phases of a typical mission profile. These are: prelaunch, ascent, on-orbit, entry and postlanding. The thermal flight data characterization section encompasses this flight data for flight to flight variations, payload effects, temperature ranges, and other variations. Discussion of the thermal environment prediction models in use by industry and various NASA Centers, and the results predicted by these models, is followed by an evaluation of the correlation between the actual flight data and the results predicted by the models. Finally, the available thermal data are evaluated from the viewpoint of the user concerned with establishing the thermal environment in the payload bay. The data deficiencies are discussed and recommendations for their elimination are presented.

Fu, J. H.↗
NASA NTRS1978

Mounting structure

A mounting platform for heat producing instruments operated in a narrow equilibrium temperature range comprises a grid-like structure with relatively large openings therein. The instruments are secured to and thermally coupled with the grid surface facing the instruments. Excess heat from the instruments is selectively radiated to the ambient through openings in the grid, the grid surfaces at these openings exhibiting low thermal emissivity and adsorptivity. The remainder of the grid is maintained at the equilibrium temperature and is covered with a thermal insulating blanket. Thus, the entire system including the platform and instruments is maintained substantially isothermal, whereby the instruments remain in fixed physical relationship to one another.

Ganssle, Eugene Robert↗
NASA NTRS2017

Thermal Control System of the ECOsystem Spaceborne Thermal Radiometer Experiment on Space Station (ECOSTRESS)

The ECOSystem Spaceborne Thermal Radiometer Experiment on Space Station (ECOSTRESS) is a thermal infrared imaging multispectral scanner science mission. ECOSTRESS is designed and built by the NASA Jet Propulsion Laboratory and will be deployed on the International Space Station (ISS). ECOSTRESS will measure the water loss from growing leaves and the evaporation from the soil by measuring the temperature of plants and will gather data on the health of the agriculture system on Earth. The ISS orbit will allow ECOSTRESS to take observations at different times during each day over the seasons to provide coverage over the contiguous United States. The ECOSTRESS Thermal Control System (TCS) consists of a combination of active and passive components to maintain the ECOSTRESS components within the allowable flight temperature (AFT) limits. The active thermal control systems include mechanical cryocoolers, heaters and a single-phase pumped fluid loop for Instrument and Payload. The focal plane detector is cooled to 65K by a pair of mechanical cryocoolers and a third mechanical cryocooler cools an intermediate cold shield to 130K. The pumped fluid loop transfers the instrument waste heat to JAXA’s Japanese Exposed Module External Facility (JEM-EF) provided external fluid loop before rejection to space from the ISS radiators. The passive TCS includes multi-layer and single layer insulations, flexible thermal links and coatings on the radiometer. This paper describes the ECOSTRESS instrument TCS architecture, instrument thermal requirements and key design drivers, the top level thermal design and analysis approach, and reports preliminary test results.

Carroll, Brian↗
NASA NTRS1991

State-of-the-art VLBI imaging - 3C345

Most VLBI images have low dynamic range because they are limited by instrumental effects such as calibration errors and poor u, v-coverage. We outline the method used to make a new image of the bright quasar 3C345 which has very high dynamic range and which is limited by the thermal noise, not instrumental errors. Both the Caltech VLBI package and the NRAO AIPS package were required to manipulate the data.

Wehrle, Ann E.↗
NASA NTRS2010

The Quest for an OCO (Orbiting Carbon Observatory) Re-Flight

The objective of the OCO (Orbiting Carbon Observatory) mission was to make the first space-based measurements of atmospheric carbon dioxide with the accuracy needed to quantify sources and sinks of this important greenhouse gas. Unfortunately, the observatory was lost as a result of a launch vehicle failure on 24 February 2009. The JPL (Jet Propulsion Laboratory) was directed to assess the options for the re-flight of the OCO instrument and recovery of the carbon-related measurement, and to understand and quantitatively assess the cost, schedule, and technical and programmatic risks of the identified options. The two most likely solutions were (1) a shared platform with the TIRS (Thermal Infrared Sensor) instrument and (2) a dedicated OSC (Orbital Sciences Corporation) LEOStar-2 spacecraft bus similar to that utilized for the original OCO mission. A joint OCO-TIRS mission study was commissioned and two specific options were examined. However, each presented technical challenges that would drive cost. It was determined that the best option was to rebuild the OCO observatory to the extent possible including another LEOStar-2 spacecraft bus. This lowest risk approach leverages the original OCO design and provides the shortest path to launch, which is targeted for no later than the February 2013 timeframe.

carbon↗
DOE OSTI2023

Development of Instrumented Advanced Test Reactor Irradiation Capsule Experiment for In-situ Thermal Conductivity Measurements of High-Density Fuels

Idaho National Laboratory (INL) is developing a first-of-a-kind leadout instrumented capsule experiment design to enable in-situ measurement capabilities in the Advanced Test Reactor (ATR) core. The Ceramic Advanced Thermal Evolution Research (CRATER) experiment supports the aLEU program objective to accelerate fuel performance irradiation testing for identifying alternative high-assay, low enriched uranium (HALEU) fuel systems. CRATER is a fueled, instrumented capsule experiment to measure in-situ temperature and thermal conductivity of ceramic fuels. Two ceramic fuel types will be used, uranium mono-nitride (UN) and uranium mono-carbide (UC), with a third metallic fuel used for comparison (UMo). The three fuel specimens will use a stainless-steel cladding. Programmatic objectives include linear heat generation rates (LHGR) of 210 ± 25 Watts per cm. and an inner clad temperature of 300-450 °C. The evolution of fuel thermal conductivity during irradiation has never been successfully measured in-situ for these systems and this experiment is designed to use advances in measurement sciences to characterize how thermal transport properties evolve while in reactor. Neutronic simulations of the experiment and its surrounding reactor environment were conducted using the Monte Carlo N-Particle Transport code (MCNP) and result in optimized fuel enrichment to meet target linear heat generation rates (LHGRs) influencing fuel temperatures, and fuel burnup requirements. Fabrication research and development (R&D) efforts are underway to produce annular right cylinder UC and UN pellets using carbothermic reduction and nitridation (or hydride-dehydride-nitride) synthesis methods, followed double-action die cold isostatic pressing.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗
NASA NTRS1986

The National Transonic Facility

The National Transonic Facility, NTF, is a high Reynolds Number facility where the increase in Reynolds Number is obtained by operating at high pressures and low temperatures. Liquid nitrogen is allowed to vaporize, making gaseous nitrogen the test medium with temperatures extending down to approximately 100 degrees Kelvin. These factors have created unique, new challenges to those developing sensors and instrumentation. Pressure vessels, thermal enclosures or elaborate temperature compensations schemes, are needed for environmental protection and special materials are needed for sensors and model fabrication. The need for a new measurement, model deformation, was also created. An extensive program to develop the unique sensors and instrumentation was initiated. The data acquisition system and systems to measure aerodynamic forces and pressures, model attitude, and model deformation, are discussed.

Holmes, H. K.↗
NASA NTRS1994

Thermal and albedo mapping of the polar regions of Mars using Viking thermal mapper observations: 2. South polar region

We present the first maps of the apparent thermal inertia and albedo of the south polar region of Mars. The observations used to create these maps were acquired by the infrared thermal mapper (IRTM) instruments on the two Viking Orbiters over a 30-day period in 1977 during the Martian late southern summer season. The maps cover the region from 60 deg S to the south pole at a spatial resolution of 1 deg of latitude, thus completing the initial thermal mapping of the entire planet. The analysis and interpretation of these maps is aided by the results of a one-dimensional radiative convective model, which is used to calculate diurnal variations in surface and atmospheric temperatures, and brightness temperatures at the top of the atmosphere for a range of assumptions concerning dust optical properties and dust optical depths. The maps show that apparent thermal inertias of bare ground regions decrease systematically from 60 deg S to the south pole. In unfrosted regions close to the south pole, apparent thermal inertias are among the lowest observed anywhere on the planet. On the south residual cap, apparent thermal inertias are very high due to the presence of CO2 frost. In most other regions of Mars, best fit apparent albedos based on thermal emission measurements are generally in good agreement with actual surface albedos based on broadband solar reflectance measurements. The one-dimensional atmospheric model calculations also predict anomalously cold brightness temperatures close to the pole during late summer, and after considering a number of alternatives, it is concluded that the net surface cooling due to atmospheric dust is the best explanation for this phenomenon. The region of lowest apparent thermal inertia close to the pole, which includes the south polar layered deposits, is interpreted to be mantled by a continuous layer of aeolian material that must be at least a few millimeters thick. The low thermal inertias mapped in the south polar region imply an absence of surface water ice deposits, which is consistent with Viking Mars atmospheric water detector (MAWD) measurements which show low atmospheric water vapor abundances throughout the summer season.

Paige, David A.↗
NASA NTRS2011

The Planned Soil Moisture Active Passive (SMAP) Mission L-Band Radar/Radiometer Instrument

The Soil Moisture Active/Passive (SMAP) mission is a NASA mission identified by the NRC 'decadal survey' to measure both soil moisture and freeze/thaw state from space. The mission will use both active radar and passive radiometer instruments at L-Band. In order to achieve a wide swath at sufficiently high resolution for both active and passive channels, an instrument architecture that uses a large rotating reflector is employed. The instrument system has completed the preliminary design review (PDR) stage, and detailed instrument design has begun. In addition to providing an overview of the instrument design, two recent design modifications are discussed: 1) The addition of active thermal control to the instrument spun side to provide a more stable, settable thermal environment for the radiometer electronics, and 2) A 'sequential transmit' strategy for the two radar polarization channels which allows a single high-power amplifier to be used.

radiometer↗
NASA NTRS2004

Thermal performance of a multi-evaporator loop heat pipe with thermal masses and thermal electrical coolers

This paper describes thermal performance of a loop heat pipe (LHP) with two evaporators and two condensers in ambient testing. Each evaporator has an outer diameter of 15mm and a length of 76mm, and has an integral compensation chamber (CC). An aluminum mass of 500 grams is attached to each evaporator to simulate the instrument mass. A thermal electric cooler (TEC) is installed on each CC to provide heating as well as cooling for CC temperature control. A flow regulator is installed in the condenser section to prevent vapor from going back to the evaporators in the event that one of condenser is fully utilized. Ammonia was used ad the working fluid. Tests conducted included start-up, power cycle, heat load sharing, sink temperature cycle, operating temperature control with TECs, and capillary limit tests. Experimental data showed that the loop could start with a heat load of less than 1OW even with added thermal masses. The loop operated stably with even and uneven evaporator heat loads, and even and uneven condenser sink temperatures. The operating temperature could be controlled within +/-0.5K of the set point temperature using either or both TECs, and the required TEC control heater power was less than 2W under most test conditions. Heat load sharing between the two evaporators was also successfully demonstrated. The loop had a heat transport capability of 120W to 140W, and could recover from a dry-out when the heat load was reduced. The 500-gram aluminum mass on each evaporator had a negligible effect on the loop operation. Existing LHPs servicing the orbiting spacecraft have a single evaporator with an outer diameter of about 25mm. Important performance characteristics demonstrated by this LHP included: 1) Operation of an LHP with 15mm diameter evaporators; 2) Robustness and reliability of an LHP with multiple evaporators and multiple condensers under various test conditions; 3) Heat load sharing among LHP evaporators; 4) Effectiveness of TECs in controlling the LHP operating temperature; and 5) Effectiveness of the flow regulator in preventing vapor from going back the evaporators.

loop heat pipes↗
NASA NTRS1985

Electron and ion density depletions measured in the STS-3 orbiter wake

The third Space Shuttle flight on Columbia carried instrumentation to measure thermal plasma density and temperature. Two separate investigations, the Plasma Diagnostics Package (PDP) and the Vehicle Charging and Potential Experiment (VCAP), carried a Langmuir Probe, and the VCAP also included a Spherical Retarding Potential Analyzer (SRPA). Only those measurements made while the PDP is in the payload bay are discussed here since the VCAP instrumentation remains in the payload bay at all times and the two measurements are compared. The wake behind a large structure (in this case the Space Shuttle Orbiter) flying through the ionospheric plasma is discussed. Much theoretical work was done regarding plasma wakes. The instrumentation on this mission gives the first data taken with a large vehicle in the ionospheric laboratory. First, the PDP Langmuir Probe and its data set will be presented, then the VCAP Langmuir Probe and SRPA with associated data. The agreement between the two data sets is discussed and then followed by some other PDP data which infers an even lower wake density.

Murphy, G. B.↗
NASA NTRS2003

WMAP Observatory Thermal Design and On-Orbit Thermal Performance

The Wilkinson Microwave Anisotropy Probe (WMAP) observatory, launched June 30, 2001, is designed to measure the cosmic microwave background radiation with unprecedented precision and accuracy while orbiting the second Lagrange point (L2). The instrument cold stage must be cooled passively to <95K, and systematic thermal variations in selected instrument components controlled to less than 0.5 mK (rms) per spin period. This paper describes the thermal design and testing of the WMAP spacecraft and instrument. Flight thermal data for key spacecraft and instrument components are presented from launch through the first year of mission operations. Effects of solar flux variation due to the Earth's elliptical orbit about the sun, surface thermo-optical property degradations, and solar flares on instrument thermal stability are discussed.

Glazer, Stuart D.↗
NASA NTRS1995

High temperature barrier coatings for refractory metals

Improvements in high temperature oxidation resistant metal coating technology will allow NASA and commercial entities to develop competitive civil space transport and communication systems. The success of investigations completed in this program will have a positive impact on broadening the technology base for high temperature materials. The work reported herein describes processes and procedures for successfully depositing coherent oxidation barrier coatings on refractory metals to prevent degradation under very severe operating environments. Application of the new technology developed is now being utilized in numerous Phase 3 applications through several prominent aerospace firms. Major achievements have included: (1) development of means to deposit thick platinum and rhodium coatings with lower stress and fewer microcracks than could be previously achieved; (2) development of processes to deposit thick adherent coatings of platinum group metals on refractory substrates that remain bonded through high temperature excursions and without need for intermediate coatings (bonding processes unique to specific refractory metals and alloys have been defined; (3) demonstration that useful alloys of refractory and platinum coatings can be made through thermal diffusion means; (4) demonstration that selected barrier coatings on refractory substrates can withstand severe oxidizing environments in the range of 1260 deg and 1760 deg C for long time periods essential to the life requirements of the hardware; and (5) successful application of the processes and procedures to prototype hardware. The results of these studies have been instrumental in improved thermal oxidation barrier coatings for the NASP propulsion system. Other Phase 3 applications currently being exploited include small uncooled thrusters for spacecraft and microsatellite maneuvering systems.

Malone, G. A.↗