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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 NTRS2000

Heat driven pulse pump

A heat driven pulse pump includes a chamber having an inlet port, an outlet port, two check valves, a wick, and a heater. The chamber may include a plurality of grooves inside wall of the chamber. When heated within the chamber, a liquid to be pumped vaporizes and creates pressure head that expels the liquid through the outlet port. As liquid separating means, the wick, disposed within the chamber, is to allow, when saturated with the liquid, the passage of only liquid being forced by the pressure head in the chamber, preventing the vapor from exiting from the chamber through the outlet port. A plurality of grooves along the inside surface wall of the chamber can sustain the liquid, which is amount enough to produce vapor for the pressure head in the chamber. With only two simple moving parts, two check valves, the heat driven pulse pump can effectively function over the long lifetimes without maintenance or replacement. For continuous flow of the liquid to be pumped a plurality of pumps may be connected in parallel.

Benner, Steve M↗
NASA NTRS1994

Thermalite Apparatus

Thermalite is portable apparatus that nondestructively determines thermal-diffusion properties of specimen. Based on heat-pulse method; it applies pulse of heat to specimen, then computes relevant thermal properties of specimen from time-resolved measurements of thermal response of specimen to heat pulse. Apparatus is self-contained and calibrated and tested with stored thermal pattern as standard. Used to assess flaw in specimen.

Heyman, Joseph S.↗
NASA NTRS1985

Determining Thermal Diffusivities of Radiating Specimens

Two heat-pulse techniques eliminate uncertainties due to radiation losses. Flashlamp supplies pulse of heat to one side of specimen. Temperatures of illumninated surface and of opposite surface monitored photometrically. In second method, pulse of electrons supplies heat to one side of specimen. Temperature of opposite side monitored by photon detector. Techniques needed because radiation heat losses in this temperature range introduce uncertainties into conventional heat-flux measurements.

Wood, C.↗
NASA NTRS2004

Temporal Treatment of a Thermal Response for Defect Depth Estimation

Transient thermography, which employs pulse surface heating of an inspected component followed by acquisition of the thermal decay stage, is gaining wider acceptance as a result of its remoteness and rapidness. Flaws in the component s material may induce a thermal contrast in surface thermograms. An important issue in transient thermography is estimating the depth of a subsurface flaw from the thermal response. This improves the quantitative ability of the thermal evaluation: from one scan it is possible to locate regions of anomalies in thickness (caused by corrosion) and estimate the implications of the flaw on the integrity of the structure. Our research focuses on thick composite aircraft components. A long square heating pulse and several minutes observation period are required to receive an adequate thermal response from such a component. Application of various time-related informative parameters of the thermal response for depth estimation is discussed. A three-dimensional finite difference model of heat propagation in solids in Cartesian coordinates is used to simulate the thermographic process. Typical physical properties of polymer graphite composites are assumed for the model.

Plotnikov, Y. A.↗
NASA NTRSDate not supplied

MEDLI2: MISP Measured Aftbody Aerothermal Environments

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars 2020 mission included two total heat flux sensors and one radiometer on the backshell to directly measure the aftbody aerothermal environments during entry into the Martian atmosphere. All three sensors successfully returned aftbody entry heating measurements. Comparisons between the total heat flux sensor measurements and DPLR/NEQAIR predictions show excellent agreement and provide confidence in the models. The radiometer measured significant radiative heating but compared to the model predictions the signal was attenuated by 50% at the end of the entry heat pulse. The loss of signal is attributed to blockage by thermal protection system (TPS) ablation product deposits on the radiometer window. Ground-based testing in the NASA Ames arc jet facilities was conducted to understand the impact of ablation product deposits on the measured radiometer signal. A discussion of the test results, how flight-like the test conditions were, and future work to further characterize the effect of TPS ablation product deposits on the radiometer performance are presented. In addition to measuring the entry heat pulse all three sensors were sensitive enough to measure solar radiation during cruise, the radiometer measured solar flux during the entry heat pulse, and the leeside total heat flux sensor picked up the descent reaction control system firings.

Ruth A Miller↗
NASA NTRS2005

Porous media matric potential and water content measurements during parabolic flight

Control of water and air in the root zone of plants remains a challenge in the microgravity environment of space. Due to limited flight opportunities, research aimed at resolving microgravity porous media fluid dynamics must often be conducted on Earth. The NASA KC-135 reduced gravity flight program offers an opportunity for Earth-based researchers to study physical processes in a variable gravity environment. The objectives of this study were to obtain measurements of water content and matric potential during the parabolic profile flown by the KC-135 aircraft. The flight profile provided 20-25 s of microgravity at the top of the parabola, while pulling 1.8 g at the bottom. The soil moisture sensors (Temperature and Moisture Acquisition System: Orbital Technologies, Madison, WI) used a heat-pulse method to indirectly estimate water content from heat dissipation. Tensiometers were constructed using a stainless steel porous cup with a pressure transducer and were used to measure the matric potential of the medium. The two types of sensors were placed at different depths in a substrate compartment filled with 1-2 mm Turface (calcined clay). The ability of the heat-pulse sensors to monitor overall changes in water content in the substrate compartment decreased with water content. Differences in measured water content data recorded at 0, 1, and 1.8 g were not significant. Tensiometer readings tracked pressure differences due to the hydrostatic force changes with variable gravity. The readings may have been affected by changes in cabin air pressure that occurred during each parabola. Tensiometer porous membrane conductivity (function of pore size) and fluid volume both influence response time. Porous media sample height and water content influence time-to-equilibrium, where shorter samples and higher water content achieve faster equilibrium. Further testing is needed to develop these sensors for space flight applications.

Parabolic Flight↗
NASA NTRSDate not supplied

Reconstruction of Thermal Protection System Aeroheating using a Green’s Function Approach

Inverse heat transfer (IHT) techniques are often used to reconstruct the surface heating conditions on spacecraft thermal protection systems (TPS) during atmospheric entry. Current IHT techniques for entry spacecraft applications, however, demand substantial computational resources, and are impractical for analyses such as uncertainty quantification and real-time health monitoring. In this paper, a Green’s function sensor fusion approach is used to reconstruct the TPS surface aeroheating conditions on experimental spaceflight and ground test systems from collocated temperature and heat flux sensors embedded in the TPS. The algorithm leverages Green’s functions to model the heat conduction within the spacecraft TPS and stabilizes the recovery of the surface heating condition using the direct heat flux sensor measurement. The algorithm is validated using arc-jet ground test data and applied to the reconstruction of the Mars 2020 backshell heating during Martian atmospheric entry. The performance of the algorithm is benchmarked against a current state-of-the-art IHT framework, FIAT_Opt. The Green’s function-based reconstruction algorithm recovers the net hot-wall heat flux absorbed by the TPS and the incident heat flux from the atmospheric entry environment in close agreement with FIAT_Opt. Notably, computation of the surface heating condition is completed in three orders of magnitude less time with the Green’s function sensor fusion approach using a consumer-grade PC, versus with FIAT_Opt running on a high performance computer cluster. The efficiency of the algorithm is leveraged to compute the uncertainty contributions of input parameters to the total uncertainty in reconstructed Mars 2020 backshell heating for the full atmospheric entry heat pulse. The sensitivity analysis uncovers that, at different times throughout the entry heat pulse, uncertainties in the TPS specific heat, thermal conductivity, and emissivity are all dominant drivers of the reconstruction uncertainty. These results demonstrate Green’s functions and sensor-fusion techniques as promising IHT approaches to reconstruct atmospheric entry environments from TPS-embedded measurements, and highlight how these techniques may give access to post-flight analyses previously hindered by the prohibitive cost of current methods.

Kenneth McAfee↗
NASA NTRSDate not supplied

Reconstruction of Thermal Protection System Aeroheating using a Green’s Function Approach

Inverse heat transfer (IHT) techniques are often used to reconstruct the surface heating conditions on spacecraft thermal protection systems (TPS) during atmospheric entry. Current IHT techniques for entry spacecraft applications, however, demand substantial computational resources, and are impractical for analyses such as uncertainty quantification and real-time health monitoring. In this paper, a Green’s function sensor fusion approach is used to reconstruct the TPS surface aeroheating conditions on experimental spaceflight and ground test systems from collocated temperature and heat flux sensors embedded in the TPS. The algorithm leverages Green’s functions to model the heat conduction within the spacecraft TPS and stabilizes the recovery of the surface heating condition using the direct heat flux sensor measurement. The algorithm is validated using arc-jet ground test data and applied to the reconstruction of the Mars 2020 backshell heating during Martian atmospheric entry. The performance of the algorithm is benchmarked against a current state-of-the-art IHT framework, FIAT_Opt. The Green’s function-based reconstruction algorithm recovers the net hot-wall heat flux absorbed by the TPS and the incident heat flux from the atmospheric entry environment in close agreement with FIAT_Opt. Notably, computation of the surface heating condition is completed in three orders of magnitude less time with the Green’s function sensor fusion approach using a consumer-grade PC, versus with FIAT_Opt running on a high performance computer cluster. The efficiency of the algorithm is leveraged to compute the uncertainty contributions of input parameters to the total uncertainty in reconstructed Mars 2020 backshell heating for the full atmospheric entry heat pulse. The sensitivity analysis uncovers that, at different times throughout the entry heat pulse, uncertainties in the TPS specific heat, thermal conductivity, and emissivity are all dominant drivers of the reconstruction uncertainty. These results demonstrate Green’s functions and sensor-fusion techniques as promising IHT approaches to reconstruct atmospheric entry environments from TPS-embedded measurements, and highlight how these techniques may give access to post-flight analyses previously hindered by the prohibitive cost of current methods.

Kenneth McAfee↗
NASA NTRS1980

Preliminary Arc-Jet tests of ablator/RSI joints in simulated space shuttle ascent and entry heating

Five ablator/RSI panels were tested in simulated Shuttle heating conditions to determine heat pulse effects on the ablator and the ablator/RSI joints. Two of the panels were subjected to both ascent and entry heating. The other three panels were subjected to entry heating only. The double-pulse heating (ascent plus entry) had no significant effect on char-layer integrity. Evaluation of the effects of entry heating on the ablator/RSI joints was difficult because, with the panel configuration used, severe pressure gradients developed within some of the joints during testing causing hot-gas flow through the joints. In joints where pressure leaks apparently did not develop, temperatures within the joints were reasonable compared to ablator and RSI back surface temperatures.

Brewer, W. D.↗
NASA NTRS1993

Simulation of ablation in Earth atmospheric entry

The process of ablation for Earth atmospheric entry is simulated using a computational approach that allows thermo-chemical nonequilibrium of the flow field and ablation gases. The heat pulse into the heat shield is modeled. The flowfield and graphite heat shield are coupled through surface mass and energy balances. The surface thermochemistry involves the oxidation of graphite and allows for catalytic recombination of diatomic oxygen. Steady-state simulations are performed on a one meter nose radius sphere at an altitude of 65/km and at freestream velocities of 8 km/s and 10 km/s. A transient simulation is performed at 65 km altitude and a freestream velocity of 10 km/s.

Keenan, James A.↗
NASA NTRS2010

Development of Thermal Protection Materials for Future Mars Entry, Descent and Landing Systems

Entry Systems will play a crucial role as NASA develops the technologies required for Human Mars Exploration. The Exploration Technology Development Program Office established the Entry, Descent and Landing (EDL) Technology Development Project to develop Thermal Protection System (TPS) materials for insertion into future Mars Entry Systems. An assessment of current entry system technologies identified significant opportunity to improve the current state of the art in thermal protection materials in order to enable landing of heavy mass (40 mT) payloads. To accomplish this goal, the EDL Project has outlined a framework to define, develop and model the thermal protection system material concepts required to allow for the human exploration of Mars via aerocapture followed by entry. Two primary classes of ablative materials are being developed: rigid and flexible. The rigid ablatives will be applied to the acreage of a 10x30 m rigid mid L/D Aeroshell to endure the dual pulse heating (peak approx.500 W/sq cm). Likewise, flexible ablative materials are being developed for 20-30 m diameter deployable aerodynamic decelerator entry systems that could endure dual pulse heating (peak aprrox.120 W/sq cm). A technology Roadmap is presented that will be used for facilitating the maturation of both the rigid and flexible ablative materials through application of decision metrics (requirements, key performance parameters, TRL definitions, and evaluation criteria) used to assess and advance the various candidate TPS material technologies.

Cassell, Alan M.↗
NASA NTRS2007

Flash Diffusivity Technique Applied to Individual Fibers

A variant of the flash diffusivity technique has been devised for determining the thermal diffusivities, and thus the thermal conductivities, of individual aligned fibers. The technique is intended especially for application to nanocomposite fibers, made from narrower fibers of polyphenylene benzobisthiazole (PBZT) and carbon nanotubes. These highly aligned nanocomposite fibers could exploit the high thermal conductivities of carbon nanotubes for thermal-management applications. In the flash diffusivity technique as practiced heretofore, one or more heat pulse(s) is (are) applied to the front face of a plate or disk material specimen and the resulting time-varying temperature on the rear face is measured. Usually, the heat pulse is generated by use of a xenon flash lamp, and the variation of temperature on the rear face is measured by use of an infrared detector. The flash energy is made large enough to produce a usefully high temperature rise on the rear face, but not so large as to significantly alter the specimen material. Once the measurement has been completed, the thermal diffusivity of the specimen is computed from the thickness of the specimen and the time dependence of the temperature variation on the rear face. Heretofore, the infrared detector used in the flash diffusivity technique has been a single-point detector, which responds to a spatial average of the thermal radiation from the rear specimen surface. Such a detector cannot distinguish among regions of differing diffusivity within the specimen. Moreover, two basic assumptions of the thermaldiffusivity technique as practiced heretofore are that the specimen is homogeneous and that heat flows one-dimensionally from the front to the rear face. These assumptions are not valid for an inhomogeneous (composite) material.

Mayeaux, Brian↗
NASA NTRS2011

An Overview of NASA's Current Materials Development Efforts for Mars EDL

Current roadmaps point to landing heavy masses (cargo, followed by manned vehicles) on Mars in the 2030's and the existing entry, descent and landing (EDL) technology will not be sufficient to facilitate such missions. In 2009 the Exploration Technology Development Program (ETDP) established the Entry, Descent and Landing Technology Development Project (EDL TDP), to be managed programmatically at Langley Research Center (LaRC) and technically a Ames Research Center (ARC). The purpose of the project is to further the technologies required to land heavy (approximately 40 metric ton) masses on Mars to facilitate exploration. The EDL TDP contains three technical elements. They are: 1) Thermal Protection Systems (TPS) development 2) Modeling and Tools (MAT) development 3) Supersonic Retropropulsion (SRP) development The primary goals of the EDL TDP TPS element is to design and develop TPS materials capable of withstanding the severe aerothermal loads associated with aerocapture and entry into the Martian atmosphere while significantly decreasing the TPS mass fraction contribution to the entry system. Significant advancements in TPS materials technology are needed in order to enable heavy mass payloads to be successfully landed on the Martian surface for robotic precursors and subsequent human exploration missions. The EDL TDP TPS element is further divided into two different TPS concepts for Mars EDL those being: 1) Rigid TPS for a mid L/D aeroshell with the capability to withstand dual pulsed heating environments as high as 500 W/square cm for aerocapture and 130 W/square cm for entry 2) Flexible TPS for a deployable aerodynamic decelerator with the capability to withstand dual pulsed heating environments as high as 120 W/square cm for aerocapture and 30 W/square cm for entry NASA, along with its vendors, has begun developing and testing materials for each of the deceleration approaches. These include multi-layer rigid ablators and flexible ablative materials. In order to model the response of these types of materials, new and improved modeling techniques will be required. This presentation will outline the types of materials that are under development and illustrate the need for advancement in modeling of ablative materials.

Beck, Robin A.↗
NASA NTRS2011

A Numerical Study of Micrometeoroids Entering Titan's Atmosphere

A study using numerical integration techniques has been performed to analyze the temperature profiles of micrometeors entering the atmosphere of Saturn s moon Titan. Due to Titan's low gravity and dense atmosphere, arriving meteoroids experience a significant cushioning effect compared to those entering the Earth's atmosphere. Temperature profiles are presented as a function of time and altitude for a number of different meteoroid sizes and entry velocities, at an entry angle of 45. Titan's micrometeoroids require several minutes to reach peak heating (ranging from 200 to 1200 K), which occurs at an altitude of about 600 km. Gentle heating may allow for gradual evaporation of volatile components over a wide range of altitudes. Computer simulations have been performed using the Cassini/Huygens atmospheric data for Titan. Keywords micrometeoroid Titan atmosphere 1 Introduction On Earth, incoming micrometeoroids (~100 m diameter) are slowed by collisions with air molecules in a relatively compact atmosphere, resulting in extremely rapid deceleration and a short heating pulse, often accompanied by brilliant meteor displays. On Titan, lower gravity leads to an atmospheric scale height that is much larger than on Earth. Thus, deceleration of meteors is less rapid and these particles undergo more gradual heating. This study uses techniques similar to those used for Earth meteoroid studies [1], exchanging Earth s planetary characteristics (e.g., mass and atmospheric profile) for those of Titan. Cassini/Huygens atmospheric data for Titan were obtained from the NASA Planetary Atmospheres Data Node [4]. The objectives of this study were 1) to model atmospheric heating of meteoroids for a range of micrometeor entry velocities for Titan, 2) to determine peak heating temperatures and rates for micrometeoroids entering Titan s atmosphere, and 3) to create a general simulation environment that can be extended to incorporate additional parameters and variables, including different atmospheric, meteoroid and planetary data. The micrometeoroid entry simulations made using Titan atmospheric data assume that, as on Earth, micrometeors are heated by collision with molecules in the atmosphere. Unlike on Earth where heating pulses last a few seconds and reach temperatures sufficient to melt silicates (> 1600 K [1]),

Templeton, M.↗
NASA NTRS2012

Analysis and Sizing for Transient Thermal Heating of Insulated Aerospace Vehicle Structures

An analytical solution was derived for the transient response of an insulated structure subjected to a simplified heat pulse. The solution is solely a function of two nondimensional parameters. Simpler functions of these two parameters were developed to approximate the maximum structural temperature over a wide range of parameter values. Techniques were developed to choose constant, effective thermal properties to represent the relevant temperature and pressure-dependent properties for the insulator and structure. A technique was also developed to map a time-varying surface temperature history to an equivalent square heat pulse. Equations were also developed for the minimum mass required to maintain the inner, unheated surface below a specified temperature. In the course of the derivation, two figures of merit were identified. Required insulation masses calculated using the approximate equation were shown to typically agree with finite element results within 10%-20% over the relevant range of parameters studied.

Blosser, Max L.↗
NASA NTRS2012

An Analytical Solution for Transient Thermal Response of an Insulated Structure

An analytical solution was derived for the transient response of an insulated aerospace vehicle structure subjected to a simplified heat pulse. This simplified problem approximates the thermal response of a thermal protection system of an atmospheric entry vehicle. The exact analytical solution is solely a function of two non-dimensional parameters. A simpler function of these two parameters was developed to approximate the maximum structural temperature over a wide range of parameter values. Techniques were developed to choose constant, effective properties to represent the relevant temperature and pressure-dependent properties for the insulator and structure. A technique was also developed to map a time-varying surface temperature history to an equivalent square heat pulse. Using these techniques, the maximum structural temperature rise was calculated using the analytical solutions and shown to typically agree with finite element simulations within 10 to 20 percent over the relevant range of parameters studied.

Blosser, Max L.↗