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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 1,189 records · Page 66
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Acoustic Database for Turbofan Engine Core-Noise Sources: Volume - I

In this program, a database of dynamic temperature and dynamic pressure measurements were acquired inside the core of a TECH977 turbofan engine to support investigations of indirect combustion noise. Dynamic temperature and pressure measurements were recorded for engine gas dynamics up to temperatures of 3100 degrees Fahrenheit and transient responses as high as 1000 hertz. These measurements were made at the entrance of the high pressure turbine (HPT) and at the entrance and exit of the low pressure turbine (LPT). Measurements were made at two circumferential clocking positions. In the combustor and inter-turbine duct (ITD), measurements were made at two axial locations to enable the exploration of time delays. The dynamic temperature measurements were made using dual thin-wire thermocouple probes. The dynamic pressure measurements were made using semi-infinite probes. Prior to the engine test, a series of bench, oven, and combustor rig tests were conducted to characterize the performance of the dual wire temperature probes and to define and characterize the data acquisition systems. A measurement solution for acquiring dynamic temperature and pressure data on the engine was defined. A suite of hardware modifications were designed to incorporate the dynamic temperature and pressure instrumentation into the TECH977 engine. In particular, a probe actuation system was developed to protect the delicate temperature probes during engine startup and transients in order to maximize sensor life. A set of temperature probes was procured and the TECH977 engine was assembled with the suite of new and modified hardware. The engine was tested at four steady state operating speeds, with repeats. Dynamic pressure and temperature data were acquired at each condition for at least one minute. At the two highest power settings, temperature data could not be obtained at the forward probe locations since the mean temperatures exceeded the capability of the probes. The temperature data were processed using software that accounts for the effects of convective and conductive heat transfer. The software was developed under previous NASA sponsored programs. Compensated temperature spectra and compensated time histories corresponding to the dynamic temperature of the gas stream were generated. Auto-spectral and cross-spectral analyses of the data were performed to investigate spectral features, acoustic circumferential mode content, signal coherence, and time delays. The dynamic temperature data exhibit a wideband and fairly flat spectral content. The temperature spectra do not change substantially with operating speed. The pressure spectra in the combustor and ITD exhibit generally similar shapes and amplitudes, making it difficult to identify any features that suggest the presence of indirect combustion noise. Cross-spectral analysis reveal a strong correlation between pressure and temperature fluctuations in the ITD, but little correlation between temperature fluctuations at the entrance of the HPT and pressure fluctuations downstream of it. Temperature fluctuations at the entrance of the low pressure turbine were an order of magnitude smaller than those at the entrance to the high pressure turbine. Time delay analysis of the temperature fluctuations in the combustor was inconclusive, perhaps due to the substantial mixing that occurs between the upstream and downstream locations. Time delay analysis of the temperature fluctuations in the ITD indicate that they convect at the mean flow speed. Analysis of the data did not reveal any convincing indications of the presence of indirect combustion noise. However, this analysis has been preliminary and additional exploration of the data is recommended including the use of more sophisticated signal processing to explore subtle issues that have been revealed but which are not yet fully understood or explained.

Aeroacoustics↗
NASA NTRS2016

Overview of the CHarring Ablator Response (CHAR) Code

An overview of the capabilities of the CHarring Ablator Response (CHAR) code is presented. CHAR is a one-, two-, and three-dimensional unstructured continuous Galerkin finite-element heat conduction and ablation solver with both direct and inverse modes. Additionally, CHAR includes a coupled linear thermoelastic solver for determination of internal stresses induced from the temperature field and surface loading. Background on the development process, governing equations, material models, discretization techniques, and numerical methods is provided. Special focus is put on the available boundary conditions including thermochemical ablation, surface-to-surface radiation exchange, and flowfield coupling. Finally, a discussion of ongoing development efforts is presented.

Amar, Adam J.↗
NASA NTRS2015

Demonstration of Hybrid Multilayer Insulation for Fixed Thickness Applications

Once on orbit, high performing insulation systems for cryogenic systems need just as good radiation (optical) properties as conduction properties. This requires the use of radiation shields with low conductivity spacers in between. By varying the height and cross-sectional area of the spacers between the radiation shields, the relative radiation and conduction heat transfers can be manipulated. However, in most systems, there is a fixed thickness or volume allocated to the insulation. In order to understand how various combinations of different multilayer insulation (MLI) systems work together and further validate thermal models of such a hybrid MLI set up, test data is needed. The MLI systems include combinations of Load Bearing MLI (LB-MLI) and traditional MLI. To further simulate the space launch vehicle case wherein both ambient pressure and vacuum environments are addressed, different cold-side thermal insulation substrates were included for select tests.

Cryogenic Fluid Management↗
NASA NTRS1961

Existence of Solutions of One-Dimensional Flow of a Real Gas

Formal solutions to the equations of motion of the one-dimensional, unsteady flow of a viscous, compressible, heat-conducting gas are presented. The relationship between the existence of these solutions and the problem of hydrodynamic stability is discussed.

Menkes, J.↗
NASA NTRS2016

Overview of the CHarring Ablator Response (CHAR) Code

An overview of the capabilities of the CHarring Ablator Response (CHAR) code is presented. CHAR is a one-, two-, and three-dimensional unstructured continuous Galerkin finite-element heat conduction and ablation solver with both direct and inverse modes. Additionally, CHAR includes a coupled linear thermoelastic solver for determination of internal stresses induced from the temperature field and surface loading. Background on the development process, governing equations, material models, discretization techniques, and numerical methods is provided. Special focus is put on the available boundary conditions including thermochemical ablation and contact interfaces, and example simulations are included. Finally, a discussion of ongoing development efforts is presented.

Amar, Adam J.↗
NASA NTRS2016

Hall Thruster Thermal Modeling and Test Data Correlation

HERMeS - Hall Effect Rocket with Magnetic Shielding. Developed through a joint effort by NASA/GRC and the Jet Propulsion Laboratory (JPL). Design goals: High power (12.5 kW) high Isp (3000 sec), high efficiency (> 60%), high throughput (10,000 kg), reduced plasma erosion and increased life (5 yrs) to support Asteroid Redirect Robotic Mission (ARRM). Further details see "Performance, Facility Pressure Effects and Stability Characterization Tests of NASAs HERMeS Thruster" by H. Kamhawi and team. Hall Thrusters (HT) inherently operate at elevated temperatures approx. 600 C (or more). Due to electric magnetic (E x B) fields used to ionize and accelerate propellant gas particles (i.e., plasma). Cooling is largely limited to radiation in vacuum environment.Thus the hardware components must withstand large start-up delta-T's. HT's are constructed of multiple materials; assorted metals, non-metals and ceramics for their required electrical and magnetic properties. To mitigate thermal stresses HT design must accommodate the differential thermal growth from a wide range of material Coef. of Thermal Expansion (CTEs). Prohibiting the use of some bolted/torqued interfaces.Commonly use spring loaded interfaces, particularly at the metal-to-ceramic interfaces to allow for slippage.However most component interfaces must also effectively conduct heat to the external surfaces for dissipation by radiation.Thus contact pressure and area are important.

Model↗
NASA NTRS2016

Electron Energetics in the Martian Dayside Ionosphere: Model Comparisons with MAVEN Data

This paper presents a study of the energetics of the dayside ionosphere of Mars using models and data from several instruments on board the Mars Atmosphere and Volatile EvolutioN spacecraft. In particular, calculated photoelectron fluxes are compared with suprathermal electron fluxes measured by the Solar Wind Electron Analyzer, and calculated electron temperatures are compared with temperatures measured by the Langmuir Probe and Waves experiment. The major heat source for the thermal electrons is Coulomb heating from the suprathermal electron population, and cooling due to collisional rotational and vibrational CO2 dominates the energy loss. The models used in this study were largely able to reproduce the observed high topside ionosphere electron temperatures (e.g., 3000 K at 300 km altitude) without using a topside heat flux when magnetic field topologies consistent with the measured magnetic field were adopted. Magnetic topology affects both suprathermal electron transport and thermal electron heat conduction. The effects of using two different solar irradiance models were also investigated. In particular, photoelectron fluxes and electron temperatures found using the Heliospheric Environment Solar Spectrum Radiation irradiance were higher than those with the Flare Irradiance Spectrum Model-Mars. The electron temperature is shown to affect the O2(+) dissociative recombination rate coefficient, which in turn affects photochemical escape of oxygen from Mars.

magnetic topology↗
NASA NTRS2016

Parametric Characterization of TES Detectors Under DC Bias

The X-ray integrated field unit (X-IFU) in European Space Agency's (ESA's) Athena mission will be the first high-resolution X-ray spectrometer in space using a large-format transition-edge sensor microcalorimeter array. Motivated by optimization of detector performance for X-IFU, we have conducted an extensive campaign of parametric characterization on transition-edge sensor (TES) detectors with nominal geometries and physical properties in order to establish sensitivity trends relative to magnetic field, dc bias on detectors, operating temperature, and to improve our understanding of detector behavior relative to its fundamental properties such as thermal conductivity, heat capacity, and transition temperature. These results were used for validation of a simple linear detector model in which a small perturbation can be introduced to one or multiple parameters to estimate the error budget for X-IFU. We will show here results of our parametric characterization of TES detectors and briefly discuss the comparison with the TES model.

X-ray Integral Field↗
NASA NTRS2017

Cooling the Origins Space Telescope

The NASA Astrophysics Division has commissioned 4 studies for consideration by the 2020 Decadal Survey to be the next flagship mission following WFIRST (Wide Field Infrared Survey Telescope). One of the four studies is the Origins Space Telescope (OST), which will cover wavelengths from 6 microns to 600 microns. To perform at the level of the zodiacal, galactic, and cosmic background, the telescope must be cooled to 4 degrees Kelvin. 4 degrees Kelvin multi-stage mechanical cryocoolers will be employed along with a multilayer sunshield/thermal shield to achieve this temperature with a manageable parasitic heat load. Current state-of-the-art cryocoolers can achieve close to 4 degrees Kelvin, providing about 50 megawatts of cooling at 4 degrees Kelvin with an input power of 500 watts. Multiple coolers at this power level will be used in parallel. These coolers also provide extra cooling power at intermediate temperature stages of 15-20 degrees Kelvin and 50-70 degrees Kelvin . This upper stage cooling will be used to limit the heat conducted to 4 degrees Kelvin . The multi-layer sunshield will limit the radiated thermal energy to the 4 degrees Kelvin volume. This paper will describe the architecture of the cryogenic system for OST along with preliminary thermal models.

cryogenics↗
NASA NTRS2017

Pressurization of a Flightweight, Liquid Hydrogen Tank: Evaporation & Condensation at a Liquid/Vapor Interface

This paper presents an analysis and simulation of evaporation and condensation at a motionless liquid/vapor interface. A 1-D model equation, emphasizing heat and mass transfer at the interface, is solved in two ways, and incorporated into a subgrid interface model within a CFD simulation. Simulation predictions are compared with experimental data from the CPST Engineering Design Unit tank, a cryogenic fluid management test tank in 1-g. The numerical challenge here is the physics of the liquid/vapor interface; pressurizing the ullage heats it by several degrees, and sets up an interfacial temperature gradient that transfers heat to the liquid phase-the rate limiting step of condensation is heat conducted through the liquid and vapor. This physics occurs in thin thermal layers O(1 mm) on either side of the interface which is resolved by the subgrid interface model. An accommodation coefficient of 1.0 is used in the simulations which is consistent with theory and measurements. This model is predictive of evaporation/condensation rates, that is, there is no parameter tuning.

Stewart, Mark E. M.↗
NASA NTRS2018

Evaluation of a Thermodynamic Ice Crystal Icing Model Using Experimental Ice Accretion Data

This paper presents the evaluation of a thermodynamic ice crystal icing model, previously presented to describe the possible mechanisms of icing within the core of a turbofan jet engine. It has been proposed that there are two types of distinct ice accretions based on a surface energy balance: freeze-dominated icing and melt-dominated icing. In the former, ice accretion occurs where a freeze fraction (0 to 1) of melted ice crystals freezes on a surface, along with the existing ice of the impinging water and ice mass. This freeze-dominated icing is characterized by having strong adhesion to the surface. In the latter, icing occurs from accumulated unmelted ice on a surface, where a melt fraction (0 to 1) dictates the amount of unmelted impinged ice. This melt-dominated icing is characterized by weakly bonded surface adhesion. The experimentally observed ice growth rates suggest that only a small fraction of the impinging ice remains on the surface, implying a mass loss mechanism such as splash, runback, bounce, or erosion. This mass loss parameter must be determined in conjunction with the fraction of freezing liquid water or fraction of melting ice on an icing surface. This loss parameter, however, along with the freeze and melt fraction, are the only experimental parameters that are currently not measured directly. Using reported icing growth rates from published ice crystal icing experiments, a methodology is proposed to determine these unknown parameters. This work takes reported ice accretion data from tests conducted by the National Aeronautics and Space Administration (NASA) in 2016 and tests NASA collaborated on with the National Research Council (NRC) of Canada in 2012 that examined the fundamental physics of ice crystal icing. Those research efforts sought to generate icing conditions representative of those that occur inside a jet engine when ingesting ice crystals. This paper presents the fundamental equations of the thermodynamic model, the methodology used to determine the aforementioned unknown icing parameters, and results from model evaluation using experimental data. In addition, this paper builds on the previously proposed model by adding a transient conduction term to explain ice growth behavior at the onset of experimental tests that was observed to be different from steady-state ice growth that occurred later in the test run.With the addition of this energy term, this becomes a quasi-steady model. A key finding from this work suggests that mass loss fractions can exceed 0.90 for steady ice growth periods. In addition, due to conductive heat fluxes when using a warmer-than-freezing airfoil, lower mass loss fraction values were calculated during the initial transient period.

Ice-Crystal Ice Accretion↗
NASA NTRS2019

A Finite Volume Procedure for Thermo-Fluid System Analysis in a Flow Network

This paper describes a finite volume procedure for network flow analysis in a thermofluid system. A flow network is defined as a group of inter-connected control volumes called “nodes” that are connected by “branches”. The mass and energy conservation equations are solved at the nodes and momentum conservation equations are solved at the branches. The flow network also includes solid nodes to account for fluid to solid heat transfer. The heat conduction equation is solved at the solid nodes in conjunction with the flow equations. The properties of a real fluid are calculated using a thermodynamic property program and used in the conservation equations. The system of equations describing the fluid-solid network is solved by a hybrid numerical method that is a combination of the Newton-Raphson and successive substitution method. This procedure has been incorporated into a general-purpose computer program, the Generalized Fluid System Simulation Program, GFSSP. This paper also presents the application and verification of the method by comparison with test data for several applications that include (a) internal flow in a rocket engine turbo-pump, (b) pressurization and loading of a cryogenic propellant tank, (c) fluid transient during sudden opening of valve for priming of an evacuated feed line, and (d) chill down of a cryogenic transfer line with phase change and two phase flows.

Majumdar, Alok↗
NASA NTRS2017

In-Flight Thermal Performance of the OCO-2 Instrument

The Orbiting Carbon Observatory-2 (OCO-2), launched on July 2, 2014, continues to operate nominally after successfully completing its primary mission of two years in space. It is hoped that, following a Senior Review and proposal phase, NASA will fund a two year extended mission, which would start in June 2017. The thermal design provides three temperature zones required by the instrument, specifically at 120 K, 267 K and 297 K. A single-stage pulse tube cryocooler provides refrigeration to three focal plane arrays to 120 K, via a high conductance flexible thermal strap. A variable conductance heat pipe (VCHP) based heat rejection system transports waste heat from the instrument, located inside the spacecraft, to the space-viewing radiators, providing tight temperature control of the spectrometer optics to 267 K, and maintains the electronics and cryocooler at 297 K. Soon after entering the A-Train on August 3, 2014, the optics and focal planes were cooled to their operating temperatures. Ice contamination of the cryogenic surfaces led to increased cryocooler loads and drove the need for two focal plane decontamination cycles between September 2014 and October 2014. A decrease in the radiometric gain of the O2 A-Band channel led to five additional focal plane decontamination cycles, between January 2015 and February 2017, and is now the primary driver of decontamination cycles. This paper provides a general overview of the thermal and cryogenic system design and reviews the inflight thermal performance for the mission.

Na-Nakornpanom, Arthur↗
NASA NTRS2019

High Temperature Superconductor Lead Assemblies for XRISM

The RESOLVE instrument on the upcoming XRISM mission is a rebuild of the SXS instrument that successfully demonstrated high resolution X-ray spectroscopy on the Hitomi mission using a microcalorimeter array at 50 mK. As with SXS, the RESOLVE instrument requires High Temperature Superconductor (HTS) lead assemblies to carry high current to the Adiabatic Demagnetization Refrigerator (ADR) while meeting a very stringent heat conductance budget. For the RESOLVE HTS lead assemblies, we have used new materials and processes to improve the stability and consistency of solder joints to the HTS material. We have also designed a new thermal strap that eliminates stress concentrations on the HTS tapes. We present the results of testing on the engineering model and flight model units, as well as the results of long term testing on prototype solder joints.

Canavan, Edgar R.↗
NASA NTRS2020

Marshall Space Flight Center Faculty Fellowship Program

This paper presents a multi-node model of autogenous pressurization of liquid nitrogen in a flight tank using the Generalized Fluid System Simulation Program (GFSSP), a general purpose flow network code developed at NASA/Marshall Space Flight Center. This study considers two different models: (a) pressurization of the tank by nitrogen gas from a supply tank and (b) pressurization of the tank by using the boil-off propellant through an IVF (Integrated Vehicle Fluids) loop. The tank pressurization model considers two different liquid fill-levels of the tank (a) 75% and (b) 45%. Heat and mass transfer between the liquid and vapor has been modeled at the liquid vapor interface. Heat transfer between wall and vapor at the ullage has been accounted for by assuming heat transfer occurs by natural convection. The model also accounts for heat leak to the tank through the insulation and metal wall by heat conduction. The predicted pressures and temperatures are compared with the measured data and good agreement has been observed. In the second case, the tank is pressurized by using the boil-off fluid that recirculates through an IVF loop. Preliminary results from this model are also reported.

N.F. Six↗
NASA NTRS2022

Effects of Increasing the Category Resolution of the Sea Ice Thickness Distribution in a Coupled Climate Model on Arctic and Antarctic Sea Ice Mean State

Many modern sea ice models used in global climate models represent the subgrid-scale heterogeneity in sea ice thickness with an ice thickness distribution (ITD), which improves model realism by representing the significant impact of the high spatial heterogeneity of sea ice thickness on thermodynamic and dynamic processes. Most models default to five thickness categories. However, little has been done to explore the effects of the resolution of this distribution (number of categories) on sea-ice feedbacks in a coupled model framework and resulting representation of the sea ice mean state. Here, we explore this using sensitivity experiments in CESM2 with the standard 5 ice thickness categories and 15 ice thickness categories. Increasing the resolution of the ITD in a run with preindustrial climate forcing results in substantially thicker Arctic sea ice year-round. Analyses show that this is a result of the ITD influence on ice strength. With 15 ITD categories, weaker ice occurs for the same average thickness, resulting in a higher fraction of ridged sea ice. In contrast, the higher resolution of thin ice categories results in enhanced heat conduction and bottom growth and leads to only somewhat increased winter Antarctic sea ice volume. The spatial resolution of the ICESat-2 satellite mission provides a new opportunity to compare model outputs with observations of seasonal evolution of the ITD in the Arctic (ICESat-2; 2018–2021). Comparisons highlight significant differences from the ITD modeled with both runs over this period, likely pointing to underlying issues contributing to the representation of average thickness.

Madison M. Smith↗
NASA NTRS2023

Origin and Evolution of Enceladus’s Tidal Dissipation

Enceladus possesses a subsurface ocean beneath a conductive ice shell. Based on shell thickness models, the estimated total conductive heat loss from Enceladus is 25–40 GW; the measured heat output from the South Polar Terrain (SPT) is 4–19 GW. The present-day SPT heat flux is of order 100 mWm -21 −, comparable to estimated paleo-heat fluxes for other regions of Enceladus. These regions have nominal ages of about 2 Ga, but the estimates are uncertain because the impactor flux in the Saturnian system may not resemble that elsewhere. Enceladus’s measured rate of orbital expansion implies a low dissipation factor Q p for Saturn, with Q p ≈3×10 -3 (neglecting the role of Dione). This value implies that Enceladus’s present-day equilibrium tidal heat production (roughly 50 GW, but with large uncertainties) is in approximate balance with its heat loss. If Q p is constant, Enceladus cannot be older than 1.5 Gyr (because otherwise it would have migrated more than is permissible). However, Saturn’s dissipation may be better described by the “resonance-locking” theory, in which case Enceladus’s orbit may have only evolved outwards by about 35% over the age of the Solar System. In the constant-Q p scenario, any ancient tidal heating events would have been too energetic to be consistent with the observations. Because resonance-locking makes capture into earlier mean-motion orbital resonances less likely, the inferred ancient heating episodes probably took place when the current orbital resonance was already established. In the resonance-locking scenario, tidal heating did not change significantly over time, allowing for a long-lived ocean and a relatively stable ice shell. If so, Enceladus is an attractive target for future exploration from a habitability standpoint.

Satellites of Saturn↗
DOE OSTI2020

Comparison between PARFUME and Bison Using the AGR-2 Irradiation Experiment

This report documents comparisons between Fuel Model (PARFUME) model predictions versus Bison for selected compacts from the second irradiation test of the Advanced Gas Reactor (AGR) program that occurred from June 2010 to October 2013 in the Advanced Test Reactor (ATR) at Idaho National Laboratory (INL). PARFUME is a fuel performance analysis and modeling code, used for evaluating gas-reactor tristructural isotropic (TRISO) coated particle fuel for prismatic, pebble bed, plate, and cylindrical type fuel geometries. PARFUME is an integrated mechanistic computer code that evaluates the thermal, mechanical, and physico-chemical behavior of TRISO coated-fuel particles and the probability for fuel failure given the particle-to-particle statistical variations in physical dimensions and material properties that arise during the fuel fabrication process. Bison is a nuclear fuel performance application built using the Multiphysics Object-Oriented Simulation Environment (MOOSE) finite element library (Permann 2020). Bison is capable of modeling multiple fuel forms in a wide variety of dimensions and geometries. It solves coupled nonlinear partial differential equations, including heat conduction, mechanics, fission product species transport etc., in a fully implicit manner. Comparisons between PARFUME and Bison were performed using four compacts from the AGR-2 experiment. Selected outputs were chosen based on their impact on the probability of the SiC layer failing. In general there was good agreement between PARFUME and Bison with the exception of predicting the gap formed between the buffer and IPyC. Further comparisons between PARFUME and Bison are planned to further develop Bison’s capabilities.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗