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At least 271 records · Page 15

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NASA NTRSDate not supplied

Flow Boiling in Microgap Coolers - Suborbital Flight Results

Flow Boiling in Microgap Coolers (FBMC) is a thermal management technique that provides embedded, on-site heat removal for power dense electronic components and systems. The dielectric coolant undergoes liquid-to-vapor phase change as it passes through channels integrated within or between devices. By facilitating direct contact between the coolant and heat-generating devices and relying on phase change, the system provides tight temperature control, hot spot mitigation, very high heat transfer coefficients, and low pumping power. Ground tests demonstrated removal of high heat fluxes (up to 500 kW/m2), very high coefficients of performance (up to 40 W of heat transferred per W of pumping power), and consistent performance across five evaporator orientations under the appropriate conditions. As a follow-on study and to better assess the role of gravity on such two-phase systems, the FBMC payload was developed and flown twice aboard the Blue Origin New Shepard space vehicle in January and May 2019. Each flight exposed the payload to weightlessness (150 seconds below 0.01 g) and shorter periods of high-g during ascent (up to 3 g) and descent (up to 5 g). The flow boiling performance was consistent independent of the acceleration levels, which provides confidence that such systems will provide consistent performance when tested on the ground, orbiting or traveling through weightlessness, and operating on the surface of the Moon, Mars, and other planetary bodies. Details of the flight results, payload development, flight environment, and planned future testing will be presented.

Flow boiling↗

NASA NTRS1999

The Effects of LOX Post Biasing on SSME Injector Wall Compatibility

An experimental investigation has been carried out to examine the effects of LOX post biasing of a shear coaxial injector on the behavior of the spray near a chamber wall. The experimental work was performed with inert propellant simulants in a high-pressure chamber. Injector flow rates and chamber pressure were designed to match the Space Shuttle Main Engine (SSME) injector gas-to-liquid density and velocity ratio at the point of propellant injection. Measurements of liquid mass flux, gas phase velocity and droplet size were made using mechanical patternation and phase Doppler interferometry techniques. The measurements revealed that the liquid mass flux distribution shifts away from the wall with increasing LOX post bias away from the wall. The shift in the liquid flux distribution was much greater than that caused by the angling of the LOX post alone. Gas velocity near the wall simultaneously increased with increasing LOX post bias away from the wall. The increase in wall side gas velocity was due to the higher fraction of gas injected on the wall side of the injector as a result of the eccentricity at the injector exit. The net result is a decrease in mixture ratio near the wall. Estimates of heat transfer and engine performance relative to the unbiased case are presented.

Strakey, P. A.↗

DOE OSTI2020

Wettability Effects on Falling Film Flow and Heat Transfer Over Horizontal Tubes in Jet Flow Mode

The performance of a falling-film heat exchanger is strongly linked to the surface characteristics and the heat transfer processes that take place over the tubes. The primary aim of this numerical study is to characterize the influence of surface wettability on the film flow behavior and its associated surface heat transfer in the jet-flow mode. Volume of fluid (VOF) based simulations are carried out for horizontal tubes with different surface wettabilities. The wettability of the tube surfaces is represented using the Kistler's dynamic contact angle model. Surface wettability effects ranging from superhydrophilic to superhydrophobic are studied by varying the equilibrium contact angle from 2 deg to 175 deg. Two different liquid mass flow rates of 0.06 and 0.18 kg/m-s corresponding to the inline and staggered jet flow modes are studied. Results are presented in terms of the liquid film thickness, the contact areas between the different phases (solid–liquid and liquid–air), and the heat transfer coefficient or Nusselt number. The resistance imposed by the increasing contact angles inhibits the extent of the liquid spreading over the tube surface, and this, in turn, influences the liquid film thickness, and the wetted area of the tube surface. A significant decrement in the heat transfer rate from the tube surfaces was observed as the equilibrium contact angle increased from 2 deg to 175 deg. The local distributions of the Nusselt number over the tube surface are strongly influenced by the flow recirculation in the liquid bulk.

Engineering↗

DOE OSTI2022

Pressure drop and heat transfer characteristics of nitrate salt and supercritical CO 2 in a diffusion-bonded heat exchanger

A lab-scale 316/L stainless steel diffusion-bonded heat exchanger was examined within a coupled high-temperature test facility consisting of an atmospheric-pressure nitrate salt loop and a supercritical carbon dioxide (sCO 2 ) loop operating between 10 and 16 MPa. Pressure loss measurements collected for flow rates up to 0.55 kg/s of Solar Salt and up to 0.6 kg/s of carbon dioxide were used to generate friction factor correlations for the respective circular and semicircular zig-zag flow passages. Heat transfer measurements were also performed over a wide range of conditions to produce Nusselt correlations for the CO 2 and nitrate salt geometries. The Nusselt number is deduced from a heat transfer resistance network and overall conductance measurements collected across the operating envelope of the coupled sCO 2 – nitrate salt test facility. The CO 2 side correlation showed excellent agreement with existing correlations in the literature. Further, the salt side, having nearly circular channels with a different zig-zag angle and hydraulic diameter, yielded a new correlation applicable to the laminar flow regime. To examine the similitude of the friction correlations, pressure drop testing with water was performed on both sides of the component as well. Finally, results from non-destructive examination are discussed to provide insights regarding inspection methods and draining performance of high-temperature liquids in compact heat exchangers.

36 MATERIALS SCIENCE↗

DOE OSTI2022

Measurement of Convective Heat Transfer Coefficients with Supercritical CO2 in Novel Additively Manufactured Helically Patterned Pin Fin Tubes Using the Wilson Plot Technique

This paper describes the measurement of convective heat transfer coefficients and friction factors for sCO2 flowing in pin-fin patterned pipes in the Heat Exchange and Experimental Testing (HEET) facility at the US DoE’s National Energy Technology Laboratory (NETL) in Morgantown, WV. The measurement procedures in the HEET rig were validated by conducting benchmark tests with smooth stainless-steel tube and comparing the results with published correlations for Nusselt number (Nu) and friction factor. Over typical Reynolds number range in sCO2, the measured Nu and friction factors were within 7% of classical correlations for smooth tube flow.The candidate pin fin patterned pipes were additively manufactured (AM) at the Oak Ridge National Laboratory. The pins were circular or elliptical in cross-section. Pin length to diameter aspect ratios were 1.33 and 2, while the pin diameter to tube diameter ratio was 0.188 and 0.125. Tests were performed for ReD varying from 6.9×104 to 2.2×105 and at conditions equivalent to the low pressure outlet (8.69 MPa, 361 K) of the low temperature recuperator (LTR) in an indirect sCO2 power cycle. The Wilson plot technique was utilized to measure the bulk heat transfer coefficients.For the better performing design (tube A, pin length to tube diameter ratio: 1.33, pin diameter to tube diameter ratio: 0.188), the local heat transfer coefficient increased by 112% relative to the Dittus-Boelter correlation at the LTR low pressure outlet. This corresponded to a 282% increase in the product of the heat transfer coefficient and the surface area. Large pressure drops across the test articles were observed.

Searle, Matthew↗

NASA NTRS1996

Development of Multiobjective Optimization Techniques for Sonic Boom Minimization

A discrete, semi-analytical sensitivity analysis procedure has been developed for calculating aerodynamic design sensitivities. The sensitivities of the flow variables and the grid coordinates are numerically calculated using direct differentiation of the respective discretized governing equations. The sensitivity analysis techniques are adapted within a parabolized Navier Stokes equations solver. Aerodynamic design sensitivities for high speed wing-body configurations are calculated using the semi-analytical sensitivity analysis procedures. Representative results obtained compare well with those obtained using the finite difference approach and establish the computational efficiency and accuracy of the semi-analytical procedures. Multidisciplinary design optimization procedures have been developed for aerospace applications namely, gas turbine blades and high speed wing-body configurations. In complex applications, the coupled optimization problems are decomposed into sublevels using multilevel decomposition techniques. In cases with multiple objective functions, formal multiobjective formulation such as the Kreisselmeier-Steinhauser function approach and the modified global criteria approach have been used. Nonlinear programming techniques for continuous design variables and a hybrid optimization technique, based on a simulated annealing algorithm, for discrete design variables have been used for solving the optimization problems. The optimization procedure for gas turbine blades improves the aerodynamic and heat transfer characteristics of the blades. The two-dimensional, blade-to-blade aerodynamic analysis is performed using a panel code. The blade heat transfer analysis is performed using an in-house developed finite element procedure. The optimization procedure yields blade shapes with significantly improved velocity and temperature distributions. The multidisciplinary design optimization procedures for high speed wing-body configurations simultaneously improve the aerodynamic, the sonic boom and the structural characteristics of the aircraft. The flow solution is obtained using a comprehensive parabolized Navier Stokes solver. Sonic boom analysis is performed using an extrapolation procedure. The aircraft wing load carrying member is modeled as either an isotropic or a composite box beam. The isotropic box beam is analyzed using thin wall theory. The composite box beam is analyzed using a finite element procedure. The developed optimization procedures yield significant improvements in all the performance criteria and provide interesting design trade-offs. The semi-analytical sensitivity analysis techniques offer significant computational savings and allow the use of comprehensive analysis procedures within design optimization studies.

Chattopadhyay, Aditi↗

NASA NTRS1993

Rover/NERVA-derived near-term nuclear propulsion

FY-92 accomplishments centered on conceptual design and analyses for 25, 50, and 75 K engines with emphasis on the 50 K engine. During the first period of performance, flow and energy balances were prepared for each of these configurations and thrust-to-weight values were estimated. A review of fuel technology and key data from the Rover/NERVA program established a baseline for proven reactor performance and areas of enhancement to meet near-term goals. Studies were performed of the criticality and temperature profiles for probable fuel and moderator loadings for the three engine sizes, with a more detailed analysis of the 50 K size. During the second period of performance, analyses of the 50 K engine continued. A chamber/nozzle contour was selected and heat transfer and fatigue analyses were performed for likely construction materials. Reactor analyses were performed to determine component radiation heating rates, reactor radiation fields, water immersion poisoning requirements, temperature limits for restartability, and a tie-tube thermal analysis. Finally, a brief assessment of key enabling technologies was made, with a view toward identifying development issues and identification of the critical path toward achieving engine qualification within 10 years.

Source record↗

DOE OSTI2021

Characterization of radiation damage in 3D printed SiC

The SiC fuel matrix for advanced gas-cooled high temperature reactors as part of the Transformational Challenge Reactor program serves as the fuel particle container structure, a barrier to fission gas release, and a heat transfer medium. Its performance is particularly important because the fuel matrix must demonstrate good structural stability and thermal behaviors. An additive manufacturing methodology combining a binder jet 3D printing process with chemical vapor infiltration (CVI) for the production of SiC was recently developed. In this study, post irradiation examination by transmission electron microscopy shows that defect accumulation within the printed particles is very similar to other forms of high-purity SiC. However, damage accumulation was not directly observed in the CVI matrix because black spot damage and dislocation loops are difficult to image within the nanoscale highly faulted CVI matrix and because interstitial defects may rapidly annihilate at the stacking faults. Therefore, electron energy loss spectroscopy (EELS) analysis was used to analyze defect swelling in both the printed particles and the CVI matrix. The EELS analysis helped reveal that the radiation-induced swelling in the CVI matrix is similar to that of the printed SiC particles. This work shows that 3D printed SiC has behavior that is comparable to SiC processed by other means and that 3D printing could serve as a suitable processing technique for high-purity SiC for nuclear applications.

36 MATERIALS SCIENCE↗

DOE OSTI2020

DUCTED ASSEMBLY STEADY STATE HEAT TRANSFER SOFTWARE

The Ducted Assembly Steady State Heat transfer software (DASSH) performs full core, steady state, subchannel temperature calculations for hexagonal, ducted assemblies. DASSH is targeted for use in the early stages of the reactor design process when assembly characteristics may not be defined or are undergoing considerable changes. DASSH provides a relatively rapid assessment of the temperature and flow distribution to inform reactor design updates and analysis.

ATZ, MILOSIVO↗

DOE OSTI2022

OTERR User Manual

OTERR (Optimization of TEst Reactor Reloading) is a software tool created to assist in the determination of optimal fuel reloading patterns for test reactors. The intended application is for the Versatile Test Reactor (VTR) program, but it provides functions that could be useful for analysis and optimization of many types of fast reactors. OTERR does not perform neutron/gamma transport, heat transfer, thermal hydraulics, or depletion calculations. Instead, it acts as a wrapper around codes that provide these capabilities, with a native genetic algorithm optimization capability. At this time, wrapping is only implemented for Argonne Reactor Computation (ARC) codes DIF3D, REBUS, and GAMSOR, and SE2-ANL. OTERR also has capabilities to facilitate input creation for DASSH, a thermal hydraulics code similar to SE2- ANL being developed for the VTR program.

21 SPECIFIC NUCLEAR REACTORS AND ASSOCIATED PLANTS↗

DOE OSTI2022

Implementation of an Orificing Optimization Algorithm in the DASSH Subchannel Analysis Code

The Ducted Assembly Steady-State Heat transfer code (DASSH) performs full-core subchannel thermal hydraulics calculations in liquid metal fast reactors. One of the applications of subchannel codes is to optimize coolant flow orificing. As a design activity, the primary task is to determine the best way to divide assemblies into groups and distribute coolant flow rates among them. This report documents an algorithm implemented in DASSH to automatically optimize coolant orificing. Over the course of multiple iterations, DASSH determines the orifice grouping and flow distribution that minimizes peak coolant, clad, or fuel temperatures across all timesteps for a user-specified number of assembly groups. The total coolant flow rate in the reactor is constrained to achieve the specified core-average outlet temperature. The flow rate to each orifice group may also be constrained by the allowable pressure drop. The distribution of coolant flow among groups is accelerated using a predictor-corrector algorithm based on interpolated results from single-assembly parametric calculations. The assembly orificing grouping is initially predicted based on assembly power but can be refined if results demonstrate that an assembly would fit better in another group. The algorithm is demonstrated with two case studies. The first is a simple model for a reactor core consisting of just fuel assemblies; the pin power distributions are specified to create a situation where the initial assembly grouping prediction is suboptimal. This example is used to describe the initial grouping, demonstrate convergence over multiple iterations, and highlight the impact of regrouping. Then, the algorithm is applied to minimize peak clad and fuel temperatures in an example sodium-cooled fast reactor, the Versatile Test Reactor. The multicycle optimization confirms prior calculations for the reference core design. The example highlights how optimizing for different peak temperatures affects the results and demonstrates the use of the pressure drop constraint to limit the maximum flow rate.

22 GENERAL STUDIES OF NUCLEAR REACTORS↗

NASA NTRS1968

Fundamental problems of injector design

The principal control of the combustion within a liquid-propellant rocket thrust chamber is provided by the injector. Only when precise control of the injection process is attained can optimum performance, dynamic stability, heat transfer compatibility and reproducibility be provided. Control must be designed into each step of the injection process with sufficient isolation provided to insure stability. In the design of the injector manifolding, the main desire is to provide mass and mixture ratio regulation across the injector face. The design of the injector orifices must be such that this mass and mixture ratio control is retained through the orifices and imparted to the droplet sprays that are produced. Furthermore, the orifice design must control the size and distribution of the droplets generated so that combustion will be completed within the chamber. Mechanical damping devices are often required to protect the early mixing zone from transverse velocities and associated displacement effects in the chamber. Only when each phase of the injection process is carefully controlled can the design aims be reached.

David T Harrje↗

NASA NTRS1973

Space vehicle integrated thermal protection/structural/meteoroid protection system, volume 1

A program was conducted to determine the merit of a combined structure/thermal meteoroid protection system for a cryogenic vehicle propulsion module. Structural concepts were evaluated to identify least weight designs. Thermal analyses determined optimum tank arrangements and insulation materials. Meteoroid penetration experiments provided data for design of protection systems. Preliminary designs were made and compared on the basis of payload capability. Thermal performance tests demonstrated heat transfer rates typical for the selected design. Meteoroid impact tests verified the protection characteristics. A mockup was made to demonstrate protection system installation. The best design found combined multilayer insulation with a truss structure vehicle body. The multilayer served as the thermal/meteoroid protection system.

Bartlett, D. H.↗

NASA NTRS1974

Similarity constraints in testing of cooled engine parts

A study is made of the effect of testing cooled parts of current and advanced gas turbine engines at the reduced temperature and pressure conditions which maintain similarity with the engine environment. Some of the problems facing the experimentalist in evaluating heat transfer and aerodynamic performance when hardware is tested at conditions other than the actual engine environment are considered. Low temperature and pressure test environments can simulate the performance of actual size prototype engine hardware within the tolerance of experimental accuracy if appropriate similarity conditions are satisfied. Failure to adhere to these similarity constraints because of test facility limitations or other reasons, can result in a number of serious errors in projecting the performance of test hardware to engine conditions.

Colladay, R. S.↗

NASA NTRS1977

Combustion effects on film cooling

The effects of: (1) a reactive environment on film cooling effectiveness, and (2) film cooling on rocket engine performance were determined experimentally in a rocket thrust chamber assembly operating with hydrogen and oxygen propellants at 300 psi chamber pressure. Tests were conducted using hydrogen, helium, and nitrogen film coolants in an instrumented, thin walled, steel thrust chamber. The film cooling, performance loss, and heat transfer coefficient data were correlated with the ALRC entrainment film cooling model which relates film coolant effectiveness and mixture ratio at the wall to the amount of mainstream gases entrained with the film coolant in a mixing layer. In addition, a comprehensive thermal analysis computer program, HOCOOL, was prepared from previously existing ALRC computer programs and analytical techniques.

Rousar, D. C.↗

NASA NTRS1976

Plasma core reactor applications

Analytical and experimental investigations were conducted to demonstrate the feasibility of fissioning uranium plasma core reactors and to characterize space and terrestrial applications for such reactors. Uranium hexafluoride fuel is injected into core cavities and confined away from the surface by argon buffer gas injected tangentially from the peripheral walls. Radiant heat transfer calculations were performed for a six-cavity reactor configuration. Axial working fluid channels are located along a fraction of each cavity peripheral wall. Results of calculations for outward-directed radiant energy fluxes corresponding to radiating temperatures of 2000 to 5000 K indicate total operating pressures from 80 to 650 atm, centerline temperatures from 6900 to 30,000 K, and total radiated powers from 25 to 2500 MW, respectively. Applications are described for this type of reactor such as (1) high-thrust, high specific impulse space propulsion, (2) highly efficient systems for generation of electricity, and (3) hydrogen or synthetic fuel production systems using the intense radiant energy fluxes.

Latham, T. S.↗

NASA NTRS1980

Advanced oxygen-hydrocarbon rocket engine study

Preliminary identification and evaluation of promising LO2/Hydrocarbon rocket engine cycles were used to produce a consistent and reliable data base for vehicle optimization and design studies. cycles G and C were chosen for design analysis. Preliminary design analysis of the heat transfer subsystem was performed to establish major technology requirements.

Obrien, C. J.↗

NASA NTRS1987

Closure models of turbulent third-order momentum and temperature fluctuations

A study is made for the development and computations of the separating and reattaching shear flows. The highlight of the study is an attempt to predict the third-moments of turbulent velocity which is responsible for the diffusion transport of the Reynolds stresses. The present computations show that the third-moments obtained by employing the law-Reynolds number model of transport equations improve the prediction of the third-moments. The modelling for scalar variables is also performed for the heat transfer computations. Since the transport equations for (u sub j theta)* and (u sub i u sub j theta) have been given in the last NASA CR, the study has been extended further to the modeling of (theta sup 2), (u sub i theta sup 2) and epsilon sub theta (dissipation rate of (theta sup 2)). The formulations are shown in this report.

Amano, R. S.↗