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NASA NTRS2008

Interface Shape Control Using Localized Heating during Bridgman Growth

Numerical calculations were performed to assess the effect of localized radial heating on the melt-crystal interface shape during vertical Bridgman growth. System parameters examined include the ampoule, melt and crystal thermal conductivities, the magnitude and width of localized heating, and the latent heat of crystallization. Concave interface shapes, typical of semiconductor systems, could be flattened or made convex with localized heating. Although localized heating caused shallower thermal gradients ahead of the interface, the magnitude of the localized heating required for convexity was less than that which resulted in a thermal inversion ahead of the interface. A convex interface shape was most readily achieved with ampoules of lower thermal conductivity. Increasing melt convection tended to flatten the interface, but the amount of radial heating required to achieve a convex interface was essentially independent of the convection intensity.

Volz, M. P.↗

NASA NTRS2010

Lightweight Heat Pipes Made from Magnesium

Magnesium has shown promise as a lighter-weight alternative to the aluminum alloys now used to make the main structural components of axially grooved heat pipes that contain ammonia as the working fluid. Magnesium heat-pipe structures can be fabricated by conventional processes that include extrusion, machining, welding, and bending. The thermal performances of magnesium heat pipes are the same as those of equal-sized aluminum heat pipes. However, by virtue of the lower mass density of magnesium, the magnesium heat pipes weigh 35 percent less. Conceived for use aboard spacecraft, magnesium heat pipes could also be attractive as heat-transfer devices in terrestrial applications in which minimization of weight is sought: examples include radio-communication equipment and laptop computers.

Rosenfeld, John N.↗

NASA NTRS2010

Modeling Of Metabolic Heat Regenerated Temperature Swing Adsorption (MTSA) Subassembly For Prototype Design

This paper describes modeling methods for the three core components of a Metabolic heat regenerated Temperature Swing Adsorption (MTSA) subassembly: a sorbent bed, a sublimation (cooling) heat exchanger (SHX), and a condensing icing (warming) heat exchanger (CIHX). The primary function of the MTSA, removing carbon dioxide from a space suit Portable Life Support System (PLSS) ventilation loop, is performed via the sorbent bed. The CIHX is used to heat the sorbent bed for desorption and to remove moisture from the ventilation loop while the SHX is alternately employed to cool the sorbent bed via sublimation of a spray of water at low pressure to prepare the reconditioned bed for the next cycle. This paper describes subsystem heat a mass transfer modeling methodologies relevant to the description of the MTSA subassembly in Thermal Desktop and SINDA/FLUINT. Several areas of particular modeling interest are discussed. In the sorbent bed, capture of the translating carbon dioxide (CO2) front and associated local energy and mass balance in both adsorbing and desorbing modes is covered. The CIHX poses particular challenges for modeling in SINDA/FLUINT as accounting for solids states in fluid submodels are not a native capability. Methods for capturing phase change and latent heat of ice as well as the transport properties across a layer of low density accreted frost are developed. This extended modeling capacity is applicable to temperatures greater than 258 K. To extend applicability to the minimum device temperature of 235 K, a method for a mapped transformation of temperatures from below the limit temperatures to some value above is given along with descriptions for associated material property transformations and the resulting impacts to total heat and mass transfer. Similar considerations are given for the SHX along with functional relationships for areal sublimation rates as limited by flow mechanics in t1he outlet duct.

Bower, Chad E.↗

NASA NTRS2010

Lunar Surface Stirling Power Systems Using Isotope Heat Sources

For many years, NASA has used the decay of plutonium-238 (Pu-238) (in the form of the General Purpose Heat Source (GPHS)) as a heat source for Radioisotope Thermoelectric Generators (RTGs), which have provided electrical power for many NASA missions. While RTGs have an impressive reliability record for the missions in which they have been used, their relatively low thermal to electric conversion efficiency and the scarcity of plutonium-238 (Pu-238) has led NASA to consider other power conversion technologies. NASA is considering returning both robotic and human missions to the lunar surface and, because of the long lunar nights (14.75 Earth days), isotope power systems are an attractive candidate to generate electrical power. NASA is currently developing the Advanced Stirling Radioisotope Generator (ASRG) as a candidate higher efficiency power system that produces greater than 160 W with two GPHS modules at the beginning of life (BOL) (~32% efficiency). The ASRG uses the same Pu-238 GPHS modules, which are used in RTG, but by coupling them to a Stirling convertor provides a four-fold reduction in the number of GPHS modules. This study considers the use of americium-241 (Am-241) as a substitute for the Pu-238 in Stirling- convertor-based Radioisotope Power Systems (RPS) for power levels from tens of watts to 5 kWe. The Am-241 is used as a substitute for the Pu-238 in GPHS modules. Depending on power level, different Stirling heat input and removal systems are modeled. It was found that substituting Am-241 GPHS modules into the ASRG reduces power output by about one-fifth while maintaining approximately the same system mass. In order to obtain the nominal 160 W of electrical output of the Pu-238 ASRG requires 10 Am-241 GPHS modules. Higher power systems require changing from conductive coupling heat input and removal from the Stirling convertor to either pumped loops or heat pipes. Liquid metal pumped loops are considered as the primary heat transportation on the hot end and water pumped loop/heat pipe radiator is considered for the heat rejection side for power levels above 1 kWe.

Schmitz, Paul C.↗

NASA NTRS2010

X-Ray Computed Tomography Inspection of the Stardust Heat Shield

The "Stardust" heat shield, composed of a PICA (Phenolic Impregnated Carbon Ablator) Thermal Protection System (TPS), bonded to a composite aeroshell, contains important features which chronicle its time in space as well as re-entry. To guide the further study of the Stardust heat shield, NASA reviewed a number of techniques for inspection of the article. The goals of the inspection were: 1) to establish the material characteristics of the shield and shield components, 2) record the dimensions of shield components and assembly as compared with the pre-flight condition, 3) provide flight infonnation for validation and verification of the FIAT ablation code and PICA material property model and 4) through the evaluation of the shield material provide input to future missions which employ similar materials. Industrial X-Ray Computed Tomography (CT) is a 3D inspection technology which can provide infonnation on material integrity, material properties (density) and dimensional measurements of the heat shield components. Computed tomographic volumetric inspections can generate a dimensionally correct, quantitatively accurate volume of the shield assembly. Because of the capabilities offered by X-ray CT, NASA chose to use this method to evaluate the Stardust heat shield. Personnel at NASA Johnson Space Center (JSC) and Lawrence Livermore National Labs (LLNL) recently performed a full scan of the Stardust heat shield using a newly installed X-ray CT system at JSC. This paper briefly discusses the technology used and then presents the following results: 1. CT scans derived dimensions and their comparisons with as-built dimensions anchored with data obtained from samples cut from the heat shield; 2. Measured density variation, char layer thickness, recession and bond line (the adhesive layer between the PICA and the aeroshell) integrity; 3. FIAT predicted recession, density and char layer profiles as well as bondline temperatures Finally suggestions are made as to future uses of this technology as a tool for non-destructively inspecting and verifying both pre and post flight heat shields.

McNamara, Karen M.↗

NASA NTRS2010

A Heat Transfer Investigation of Liquid and Two-Phase Methane

A heat transfer investigation was conducted for liquid and two-phase methane. The tests were conducted at the NASA Glenn Research Center Heated Tube Facility (HTF) using resistively heated tube sections to simulate conditions encountered in regeneratively cooled rocket engines. This testing is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. Nontoxic propellants, such as liquid oxygen/liquid methane (LO2/LCH4), offer potential benefits in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications. Regeneratively cooled thrust chambers are one solution for high performance, robust LO2/LCH4 engines, but cooling data on methane is limited. Several test runs were conducted using three different diameter Inconel 600 tubes, with nominal inner diameters of 0.0225-, 0.054-, and 0.075-in. The mass flow rate was varied from 0.005 to 0.07 lbm/sec. As the current focus of the PCAD project is on pressure fed engines for LO2/LCH4, the average test section outlet pressures were targeted to be 200 psia or 500 psia. The heat flux was incrementally increased for each test condition while the test section wall temperatures were monitored. A maximum average heat flux of 6.2 Btu/in.2 sec was achieved and, at times, the temperatures of the test sections reached in excess of 1800 R. The primary objective of the tests was to produce heat transfer correlations for methane in the liquid and two-phase regime. For two-phase flow testing, the critical heat flux values were determined where the fluid transitions from nucleate boiling to film boiling. A secondary goal of the testing was to measure system pressure drops in the two-phase regime.

VanNoord, Jonathan↗

NASA NTRS2011

Thermal Vacuum Testing of a Proto-flight Miniature Loop Heat Pipe with Two Evaporators and Two Condensers

This paper describes thermal vacuum testing of a proto-flight miniature loop heat pipe (MLHP) with two evaporators and two condensers designed for future small systems applications requiring low mass, low power and compactness. Each evaporator contains a wick with an outer diameter of 6.35 mm, and each has its own integral compensation chamber (CC). Miniaturization of the loop components reduces the volume and mass of the thermal system. Multiple evaporators provide flexibility for placement of instruments that need to be maintained at the same temperature, and facilitate heat load sharing among instruments, reducing the auxiliary heater power requirement. A flow regulator is used to regulate heat dissipations between the two condensers, allowing flexible placement of radiators on the spacecraft. A thermoelectric converter (TEC) is attached to each CC for control of the operating temperature and enhancement of start-up success. Tests performed include start-up, power cycle, sink temperature cycle, high power and low power operation, heat load sharing, and operating temperature control. The proto-flight MLHP demonstrated excellent performance in the thermal vacuum test. The loop started successfully and operated stably under various evaporator heat loads and condenser sink temperatures. The TECs were able to maintain the loop operating temperature within b1K of the desired set point temperature at all power levels and all sink temperatures. The un-powered evaporator would automatically share heat from the other powered evaporator. The flow regulator was able to regulate the heat dissipation among the radiators and prevent vapor from flowing into the liquid line.

Ku, Jentung↗

NASA NTRS2008

Calculations of Solar Shortwave Heating Rates due to Black Carbon and Ozone Absorption Using in Situ Measurements

Results for the solar heating rates in ambient air due to absorption by black-carbon (BC) containing particles and ozone are presented as calculated from airborne observations made in the tropical tropopause layer (TTL) in January-February 2006. The method uses airborne in situ observations of BC particles, ozone and actinic flux. Total BC mass is obtained along the flight track by summing the masses of individually detected BC particles in the range 90 to 600-nm volume-equivalent diameter, which includes most of the BC mass. Ozone mixing ratios and upwelling and partial downwelling solar actinic fluxes were measured concurrently with BC mass. Two estimates used for the BC wavelength-dependent absorption cross section yielded similar heating rates. For mean altitudes of 16.5, 17.5, and 18.5 km (0.5 km) in the tropics, average BC heating rates were near 0.0002 K/d. Observed BC coatings on individual particles approximately double derived BC heating rates. Ozone heating rates exceeded BC heating rates by approximately a factor of 100 on average and at least a factor of 4, suggesting that BC heating rates in this region are negligible in comparison.

Gao, R. S.↗

NASA NTRS2011

Base Heating Sensitivity Study for a 4-Cluster Rocket Motor Configuration in Supersonic Freestream

In support of launch vehicle base heating and pressure prediction efforts using the Loci-CHEM Navier-Stokes computational fluid dynamics solver, 35 numerical simulations of the NASA TND-1093 wind tunnel test have been modeled and analyzed. This test article is composed of four JP-4/LOX 500 lbf rocket motors exhausting into a Mach 2 - 3.5 wind tunnel at various ambient pressure conditions. These water-cooled motors are attached to a base plate of a standard missile forebody. We explore the base heating profiles for fully coupled finite-rate chemistry simulations, one-way coupled RAMP (Reacting And Multiphase Program using Method of Characteristics)-BLIMPJ (Boundary Layer Integral Matrix Program - Jet Version) derived solutions and variable and constant specific heat ratio frozen flow simulations. Variations in turbulence models, temperature boundary conditions and thermodynamic properties of the plume have been investigated at two ambient pressure conditions: 255 lb/sq ft (simulated low altitude) and 35 lb/sq ft (simulated high altitude). It is observed that the convective base heat flux and base temperature are most sensitive to the nozzle inner wall thermal boundary layer profile which is dependent on the wall temperature, boundary layer s specific energy and chemical reactions. Recovery shock dynamics and afterburning significantly influences convective base heating. Turbulence models and external nozzle wall thermal boundary layer profiles show less sensitivity to base heating characteristics. Base heating rates are validated for the highest fidelity solutions which show an agreement within +/-10% with respect to test data.

Mehta, Manish↗

NASA NTRS2010

Relating Convective and Stratiform Rain to Latent Heating

The relationship among surface rainfall, its intensity, and its associated stratiform amount is established by examining observed precipitation data from the Tropical Rainfall Measuring Mission (TRMM) Precipitation Radar (PR). The results show that for moderate-high stratiform fractions, rain probabilities are strongly skewed toward light rain intensities. For convective-type rain, the peak probability of occurrence shifts to higher intensities but is still significantly skewed toward weaker rain rates. The main differences between the distributions for oceanic and continental rain are for heavily convective rain. The peak occurrence, as well as the tail of the distribution containing the extreme events, is shifted to higher intensities for continental rain. For rainy areas sampled at 0.58 horizontal resolution, the occurrence of conditional rain rates over 100 mm/day is significantly higher over land. Distributions of rain intensity versus stratiform fraction for simulated precipitation data obtained from cloud-resolving model (CRM) simulations are quite similar to those from the satellite, providing a basis for mapping simulated cloud quantities to the satellite observations. An improved convective-stratiform heating (CSH) algorithm is developed based on two sources of information: gridded rainfall quantities (i.e., the conditional intensity and the stratiform fraction) observed from the TRMM PR and synthetic cloud process data (i.e., latent heating, eddy heat flux convergence, and radiative heating/cooling) obtained from CRM simulations of convective cloud systems. The new CSH algorithm-derived heating has a noticeably different heating structure over both ocean and land regions compared to the previous CSH algorithm. Major differences between the new and old algorithms include a significant increase in the amount of low- and midlevel heating, a downward emphasis in the level of maximum cloud heating by about 1 km, and a larger variance between land and ocean in the new CSH algorithm.

Tao, Wei-Kuo↗

NASA NTRS2011

Ascent Heating Thermal Analysis on Spacecraft Adaptor Fairings

When the Crew Exploration Vehicle (CEV) is launched, the spacecraft adaptor (SA) fairings that cover the CEV service module (SM) are exposed to aero heating. Thermal analysis is performed to compute the fairing temperatures and to investigate whether the temperatures are within the material limits for nominal ascent aeroheating case. The ascent heating is analyzed by using computational fluid dynamics (CFD) and engineering codes at Marshall Space Flight Center. The aeroheating environment data used for this work is known as Thermal Environment 3 (TE3) heating data. One of the major concerns is with the SA fairings covering the CEV SM and the SM/crew launch vehicle (CLV) flange interface. The TE3 heating rate is a function of time, wall temperature, and the spatial locations. The implementation of the TE3 heating rate as boundary conditions in the thermal analysis becomes challenging. The ascent heating thermal analysis on SA fairings and SM/CLV flange interface are performed using two commercial software packages: Cullimore & Ring (C&R) Thermal Desktop (TD) 5.1 and MSC Patran 2007r1 b. TD is the pre-and post-processor for SINDA, which is a finite-difference-based solver. In TD, the geometry is built and meshed, the boundary conditions are defined, and then SINDA is used to compute temperatures. MSC Pthermal is a finite-element- based thermal solver. MSC Patran is the pre- and post-processor for Pthermal. Regarding the boundary conditions, the convection, contact resistance, and heat load can be imposed in different ways in both programs. These two software packages are used to build the thermal model for the same analysis to validate each other and show the differences in the modeling details.

Wang, Xiao Yen↗

NASA NTRS2011

Conjugate Compressible Fluid Flow and Heat Transfer in Ducts

A computational approach to modeling transient, compressible fluid flow with heat transfer in long, narrow ducts is presented. The primary application of the model is for analyzing fluid flow and heat transfer in solid propellant rocket motor nozzle joints during motor start-up, but the approach is relevant to a wide range of analyses involving rapid pressurization and filling of ducts. Fluid flow is modeled through solution of the spatially one-dimensional, transient Euler equations. Source terms are included in the governing equations to account for the effects of wall friction and heat transfer. The equation solver is fully-implicit, thus providing greater flexibility than an explicit solver. This approach allows for resolution of pressure wave effects on the flow as well as for fast calculation of the steady-state solution when a quasi-steady approach is sufficient. Solution of the one-dimensional Euler equations with source terms significantly reduces computational run times compared to general purpose computational fluid dynamics packages solving the Navier-Stokes equations with resolved boundary layers. In addition, conjugate heat transfer is more readily implemented using the approach described in this paper than with most general purpose computational fluid dynamics packages. The compressible flow code has been integrated with a transient heat transfer solver to analyze heat transfer between the fluid and surrounding structure. Conjugate fluid flow and heat transfer solutions are presented. The author is unaware of any previous work available in the open literature which uses the same approach described in this paper.

Cross, M. F.↗

NASA NTRS2012

Heated-Atmosphere Airship for the Titan Environment: Thermal Analysis

Future exploration of Saturn's moon Titan can be carried out by airships. Several lighter-than-atmosphere gas airships and passive drifting heated-atmosphere balloon designs have been studied, but a heated-atmosphere airship could combine the best characteristics of both. This work analyses the thermal design of such a heated-atmosphere vehicle, and compares the result with a lighter-than-atmosphere (hydrogen) airship design. A design tool was created to enable iteration through different design parameters of a heated-atmosphere airship (diameter, number of layers, and insulating gas pocket thicknesses) and evaluate the feasibility of the resulting airship. A baseline heated-atmosphere airship was designed to have a diameter of 6 m (outer diameter of 6.2 m), three-layers of material, and an insulating gas pocket thickness of 0.05 m between each layer. The heated-atmosphere airship has a mass of 161.9 kg. A similar mission making use of a hydrogen-filled airship would require a diameter of 4.3 m and a mass of about 200 kg. For a long-duration mission, the heated-atmosphere airship appears better suited. However, for a mission lifetime under 180 days, the less complex hydrogen airship would likely be a better option.

Heller, R. S.↗

NASA NTRS2012

James Webb Space Telescope Integrated Science Instrument Module Calibration and Verification of High-Accuracy Instrumentation to Measure Heat Flow in Cryogenic Testing

The James Webb Space Telescope (JWST) is an upcoming flagship observatory mission scheduled to be launched in 2018. Three of the four science instruments are passively cooled to their operational temperature range of 36K to 40K, and the fourth instrument is actively cooled to its operational temperature of approximately 6K. The requirement for multiple thermal zoned results in the instruments being thermally connected to five external radiators via individual high purity aluminum heat straps. Thermal-vacuum and thermal balance testing of the flight instruments at the Integrated Science Instrument Module (ISIM) element level will take place within a newly constructed shroud cooled by gaseous helium inside Goddard Space Flight Center's (GSFC) Space environment Simulator (SES). The flight external radiators are not available during ISIM-level thermal vacuum/thermal testing, so they will be replaced in test with stable and adjustable thermal boundaries with identical physical interfaces to the flight radiators. Those boundaries are provided by specially designed test hardware which also measures the heat flow within each of the five heat straps to an accuracy of less than 2 mW, which is less than 5% of the minimum predicted heat flow values. Measurement of the heat loads to this accuracy is essential to ISIM thermal model correlation, since thermal models are more accurately correlated when temperature data is supplemented by accurate knowledge of heat flows. It also provides direct verification by test of several high-level thermal requirements. Devices that measure heat flow in this manner have historically been referred to a "Q-meters". Perhaps the most important feature of the design of the JWST Q-meters is that it does not depend on the absolute accuracy of its temperature sensors, but rather on knowledge of precise heater power required to maintain a constant temperature difference between sensors on two stages, for which a table is empirically developed during a calibration campaign in a small chamber at GSFC. This paper provides a brief review of Q-meter design, and discusses the Q-meter calibration procedure including calibration chamber modifications and accommodations, handling of differing conditions between calibration and usage, the calibration process itself, and the results of the tests used to determine if the calibration is successful.

Comber, Brian↗

NASA NTRS2013

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames. Nuclear Thermal Rockets (NTR) are capable of producing a high specific impulse by employing heat produced by a fission reactor to heat and therefore accelerate hydrogen through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high‐temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements due to large thermal gradients; therefore, high‐melting‐point ceramics‐metallic matrix composites (cermets) are one of the fuels under consideration as part of the Nuclear Cryogenic Propulsion Stage (NCPS) Advance Exploration System (AES) technology project at the Marshall Space Flight Center. The purpose of testing and analytical modeling is to determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures and obtain data to assess the properties of the non‐nuclear support materials. The fission process and the resulting heating performance are well known and do not require that active fissile material to be integrated in this testing. A small‐scale test bed; Compact Fuel Element Environmental Tester (CFEET), designed to heat fuel element samples via induction heating and expose samples to hydrogen is being developed at MSFC to assist in optimal material and manufacturing process selection without utilizing fissile material. This paper details the analytical approach to help design and optimize the test bed using COMSOL Multiphysics for predicting thermal gradients induced by electromagnetic heating (Induction heating) and Thermal Desktop for radiation calculations.

Gomez, C. F.↗

NASA NTRS2013

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

Gomez, Carlos F.↗

NASA NTRS2014

Heat Pipe Planets

When volcanism dominates heat transport, a terrestrial body enters a heat-pipe mode, in which hot magma moves through the lithosphere in narrow channels. Even at high heat flow, a heat-pipe planet develops a thick, cold, downwards-advecting lithosphere dominated by (ultra-)mafic flows and contractional deformation at the surface. Heat-pipes are an important feature of terrestrial planets at high heat flow, as illustrated by Io. Evidence for their operation early in Earth's history suggests that all terrestrial bodies should experience an episode of heat-pipe cooling early in their histories.

Moore, William B.↗

NASA NTRS2014

Visualization of Capsule Reentry Vehicle Heat Shield Ablation Using Naphthalene PLIF

The Orion Multi-Purpose Crew Vehicle (MPCV) will use an ablative heat shield and improved understanding of the ablation process would be beneficial for design purposes. Given that ablation is a multi-physics process involving heat and mass transfer, codes aiming to predict heat shield ablation are in need of experimental data pertaining to the turbulent transport of ablation products for validation. At The University of Texas at Austin, a technique is being developed that uses planar laser-induced fluorescence (PLIF) of a low-temperature sublimating ablator (naphthalene) to visualize the transport of ablation products in a supersonic flow. Since ablation at reentry temperatures can be difficult to recreate in a laboratory setting it is desirable to create a limited physics problem and simulate the ablation process at relatively low temperature conditions using naphthalene. A scaled Orion MPCV model with a solid naphthalene heat shield has been tested in a Mach 5 wind tunnel at various angles of attack in the current work. PLIF imaging reveals the distribution of the ablation products as they are transported into the heat-shield boundary layer and over the capsule shoulders into the separated shear layer and backshell recirculation region. Visualizations of the capsule shear layer using both naphthalene PLIF and Schlieren imaging compared favorably. High concentrations of naphthalene in the capsule separated flow region, intermittent turbulent structures on the heat shield surface, and interesting details of the capsule shear layer structure were observed using the naphthalene PLIF technique. The capsule shear layer was also shown to generally appear to be more turbulent at lower angles of attack. Furthermore, the PLIF signal increased steadily over the course of a run indicating that during a wind tunnel run the model heated up and the rate of naphthalene ablation increased. The shear layer showed increasing signs of turbulence over the course of a wind tunnel run as well, likely because of the combination of increased surface roughness and surface blowing rate. PLIF imaging also detected regions with a relatively low concentration of naphthalene in the capsule backshell recirculation region that are most likely the result of cross-flow-induced vortices on the capsule afterbody.

Combs, Christopher S.↗