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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 793 records · Page 44
All sourcesNASA NTRSDOE OSTI
DOE OSTI2024

Probabilistic locked mode predictor in the presence of a resistive wall and finite island saturation in tokamaks

We present a framework for estimating the probability of locking to an error field in a rotating tokamak plasma. This leverages machine learning methods trained on data from a mode-locking model, including an error field, resistive magnetohydrodynamics modeling of the plasma, a resistive wall, and an external vacuum region, leading to a fifth-order ordinary differential equation (ODE) system. It is an extension of the model without a resistive wall introduced by Akçay et al. [Phys. Plasmas 28, 082106 (2021)]. Tearing mode saturation by a finite island width is also modeled. We vary three pairs of control parameters in our studies: the momentum source plus either the error field, the tearing stability index, or the island saturation term. The order parameters are the time-asymptotic values of the five ODE variables. Normalization of them reduces the system to 2D and facilitates the classification into locked (L) or unlocked (U) states, as illustrated by Akçay et al., [Phys. Plasmas 28, 082106 (2021)]. This classification splits the control space into three regions: L̂, with only L states; Û, with only U states; and a hysteresis (hysteretic) region Ĥ, with both L and U states. In regions L̂ and Û, the cubic equation of torque balance yields one real root. Region Ĥ has three roots, allowing bifurcations between the L and U states. The classification of the ODE solutions into L/U is used to estimate the locking probability, conditional on the pair of the control parameters, using a neural network. We also explore estimating the locking probability for a sparse dataset, using a transfer learning method based on a dense model dataset.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1995

Resolution of the buoyancy in the 8-foot high temperature tunnel combustor

Currently, the 8-Foot High Temperature Tunnel (8-Ft. HTT) combustor produces a good profile at only one point (2000 psia and 3650 R with oxygen enrichment). Air is enriched with oxygen (liquid) so that the combustor product gas will contain the volumetric amount of oxygen normally found in air. The oxygen enriched air has a large fraction that is not reacted and flows through the outer periphery of the fuel injector. This ring of cold air in addition to the relatively cold walls of the combustor set up buoyancy forces that produce a segregation of relatively cool gases at the bottom of the combustor exit. The basic problem is to produce a test gas that has uniform properties at all combustor conditions. The combustor temperature may be as high as 3700 R or as low as 2000 R. Combustor pressures can be as high as 3500 psia (no oxygen enrichment) and as low as 600 psia. The segregation is most severe with oxygen enriched air, since its temperature is lower and its density is high. The combustor is lined with nickel 201 and can be operated at about 1600 R maximum. A global mixing process is desired that produces an acceptable profile of temperature, species, and velocity at the exit of the combustor. The ultimate goal is a temperature profile with about 100 R variance and about 2 percent variance in oxygen. The exit total temperature must not be lowered significantly by the mixing apparatus or mechanisms employed. If immersed bodies are used, they must also be kept very hot. All combustor wall modifications must be able to survive the heat and structural conditions of the varied operating conditions. Our approach to resolving this issue is being conducted in three stages: (1) Consider mixing exclusively, (2) Resolve the heat transfer concerns resulting from the chosen mixing strategy, and (3) Solve the material and structural problems resulting from stages (1) and (2). Since the 8-Ft. HTT is unavailable for experimentation, the study is conducted exclusively with computational fluid dynamic (CPD) codes (Fluent/Uns and Rampant 3.1) using unstructured grid through body fitted coordinates. Both CFD codes are general purpose Navier-Stokes solution packages that can solve integral conservation equations for conservation of mass, momentum and energy. The governing equations are discretized using a control-volume finite-element method on unstructured triangular 2-D grids. In the interest of time, a 3-D tetrahedral grid was used to check the 2-D results on one mixing strategy and the 2-D results were confirmed. Preliminary results indicate that excellent mixing can be achieved with a body placed in the center of the flowing hot fluid with a minor modification to the combustor wall similar to a model positioned in the test section of a wind tunnel. The concept here, is to create longitudinal vortices strong enough to bring both fluids into intimate contact with each other near the wall where the cold fluid resides. However, there may be a trade-off in the choice of mixing strategy, heat transfer and structural requirements. Currently, we are examining the more promising geometries for heat transfer concerns and developing strategies for the material of construction of the center body.

Loney, Norman W.↗
DOE OSTI2022

Reassessing energy deposition for the ITER 5 MA vertical displacement event with an improved DINA model

The beryllium (Be) main chamber wall interaction during a 5 MA/1.8 T upward, unmitigated VDE scenario, first analysed in [J. Coburn et al., Phys. Scr. T171 (2020) 014076] for ITER, has been re-evaluated using the latest energy deposition analysis software. Updates to the DINA disruption model are summarized, including an improved numerical convergence for the 0D power balance, limitations on the safety factor within the plasma core, and the choice to maintain a constant plasma +halo poloidal cross-section. Such updates result in a broad halo region and higher radiated power fractions compared to previous models. The new scenario lasts for ~75 ms and deposits ~29 MJ of energy, with the radial distribution of parallel heat flux q‖(r)resembling an exponential falloff with an effective λ_q=75-198 mm. A maximum halo width w_h of 0.52 m at the outboard midplane is observed. SMITER field line tracing and energy deposition simulations calculate a q_{⊥,max} of ~83 MW/m^2 on the upper first wall panels (FWP). Heat transfer calculations with the MEMOS-U code show that the FWP surface temperature reaches ~1000 K, well below the Be melt threshold. Variations of this 5 MA scenario with Be im-purity densities from 0 to 3∙10^19 m^-3 also remain below the melt threshold despite differences in energy deposition and duration. These results are in contrast to the early study which predicted melt damage to the first wall [J. Coburn et al., Phys. Scr. T171 (2020) 014076], and emphasize the importance of accurate models for the halo width w_h and the heat flux distribution q‖(r)within that halo width. The 2020 halo model in DINA has been compared with halo current experiments on COMPASS, JET, and Alcator C-Mod, and the preliminary results build confidence in the broad halo width predictions. Results for the 5 MA VDE are compared with those for a 15 MA equivalent, generated using the new DINA model. At the higher current, significant melting of the upper FWP is to be expected.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1991

Proper orthogonal decomposition based turbulence modeling

The proper orthogonal decomposition is utilized to provide one mode estimates of various turbulence quantities in the near-wall region of the turbulent boundary layer. The experimental data of Herzog (1986) and the direct numerical simulation (DNS) results of Zang (1991) proved the two-point correlation tensor. The correlation tensor is then used as the kernel in the integral eigenvalue problem from which the proper orthogonal modes are obtained. The one mode estimates of the turbulent kinetic energy, dissipation, and triple moments all exhibit the proper asymtotic behavior near the wall. As a validation test, a model for the triple moments is developed and its near-wall behavior is studied. This information can then be used to obtain an estimate of the damping function near the wall, which is then compared to known existing behavior. Since this is an initial investigation, the streamwise and spanwise dependence is surpressed and only the wall normal direction (y+) behavior of various turbulence statistics near the wall is studied. Before examining one mode estimates of the triple moments and the dissipation (on which the model is based), various second order moments are examined. If these do not exhibit the proper behavior near the wall (as discussed by Speziale et al 1990) it does not seem reasonable to expect the one mode estimate of the triple moment and dissipation to have the right behavior. Only the results obtained using the experimental data of Herzog are discussed since the results from the simulations of Zang exhibit similar behavior as discussed above.

Glauser, Mark N.↗
NASA NTRS2020

Computational Simulations of a Mach 0.745 Transonic Truss-Braced Wing Design

A joint effort between the NASA Ames and Langley Research Centers was undertaken to analyze the Mach 0.745 variant of the Boeing Transonic Truss-Braced Wing (TTBW) Design. Two different flow solvers, LAVA and USM3D, were used to predict the TTBW flight performance. Sensitivity studies related to mesh resolution and numerical schemes were conducted to define best practices for this type of geometry and flow regime. Validation efforts compared the numerical simulation results of various modeling methods against experimental data taken from the NASA Ames 11-foot Unitary Wind Tunnel experimental data. The fidelity of the computational representation of the wind tunnel experiment, such as utilizing a porous wall boundary condition to model the ventilated test section, was varied to examine how different tunnel effects influence CFD predictions. LAVA and USM3D results both show an approximate 0.5 angle of attack shift from experimental lift curve data. This drove an investigation that revealed that the trailing edge of the experimental model was rounded in comparison to the CAD model, due to manufacturing tolerances, which had not been accounted for in the initial simulations of the experiment. Simulating the TTBW with an approximation of this rounded trailing-edge reduces error by approximately 60%. An accurate representation of the tested TTBW geometry, ideally including any wing twists and deflections experienced during the test under various loading conditions, will be necessary for proper validation of the CFD.

Maldonado, Daniel↗
NASA NTRS2020

Computational Simulations of a Mach 0.745 Transonic Truss-Braced Wing Design

A joint effort between the NASA Ames and Langley Research Centers was undertaken to analyze the Mach 0.745 variant of the Boeing Transonic Truss-Braced Wing (TTBW) Design. Two different flow solvers, LAVA and USM3D, were used to predict the TTBW flight performance. Sensitivity studies related to mesh resolution and numerical schemes were conducted to define best practices for this type of geometry and flow regime. Validation efforts compared the numerical simulation results of various modeling methods against experimental data taken from the NASA Ames 11-foot Unitary Wind Tunnel experimental data. The fidelity of the computational representation of the wind tunnel experiment, such as utilizing a porous wall boundary condition to model the ventilated test section, was varied to examine how different tunnel effects influence CFD predictions. LAVA and USM3D results both show an approximate 0.5o angle of attack shift from experimental lift curve data. This drove an investigation that revealed that the trailing edge of the experimental model was rounded in comparison to the CAD model, due to manufacturing tolerances, which had not been accounted for in the initial simulations of the experiment. Simulating the TTBW with an approximation of this rounded trailing-edge reduces error by approximately 60%. An accurate representation of the tested TTBW geometry, ideally including any wing twists and deflections experienced during the test under various loading conditions, will be necessary for more thorough validation of the CFD.

Maldonado, Daniel↗
NASA NTRS1994

Numerical Analysis of an Impinging Jet Reactor for the CVD and Gas-Phase Nucleation of Titania

We model a cold-wall atmospheric pressure impinging jet reactor to study the CVD and gas-phase nucleation of TiO2 from a titanium tetra-iso-propoxide (TTIP)/oxygen dilute source gas mixture in nitrogen. The mathematical model uses the computational code FIDAP and complements our recent asymptotic theory for high activation energy gas-phase reactions in thin chemically reacting sublayers. The numerical predictions highlight deviations from ideality in various regions inside the experimental reactor. Model predictions of deposition rates and the onset of gas-phase nucleation compare favorably with experiments. Although variable property effects on deposition rates are not significant (approximately 11 percent at 1000 K), the reduction rates due to Soret transport is substantial (approximately 75 percent at 1000 K).

Gokoglu, Suleyman A.↗
NASA NTRS2016

Comparison of Experimental Surface and Flow Field Measurements to Computational Results of the Juncture Flow Model

Wing-body juncture flow fields on commercial aircraft configurations are challenging to compute accurately. The NASA Advanced Air Vehicle Program's juncture flow committee is designing an experiment to provide data to improve Computational Fluid Dynamics (CFD) modeling in the juncture flow region. Preliminary design of the model was done using CFD, yet CFD tends to over-predict the separation in the juncture flow region. Risk reduction wind tunnel tests were requisitioned by the committee to obtain a better understanding of the flow characteristics of the designed models. NASA Ames Research Center's Fluid Mechanics Lab performed one of the risk reduction tests. The results of one case, accompanied by CFD simulations, are presented in this paper. Experimental results suggest the wall mounted wind tunnel model produces a thicker boundary layer on the fuselage than the CFD predictions, resulting in a larger wing horseshoe vortex suppressing the side of body separation in the juncture flow region. Compared to experimental results, CFD predicts a thinner boundary layer on the fuselage generates a weaker wing horseshoe vortex resulting in a larger side of body separation.

Roozeboom, Nettie H.↗
NASA NTRS2016

Comparison of Experimental Surface and Flow Field Measurements to Computational Results of the Juncture Flow Model

Wing-body juncture flow fields on commercial aircraft configurations are challenging to compute accurately. The NASA Advanced Air Vehicle Program's juncture flow committee is designing an experiment to provide data to improve Computational Fluid Dynamics (CFD) modeling in the juncture flow region. Preliminary design of the model was done using CFD, yet CFD tends to over-predict the separation in the juncture flow region. Risk reduction wind tunnel tests were requisitioned by the committee to obtain a better understanding of the flow characteristics of the designed models. NASA Ames Research Center's Fluid Mechanics Lab performed one of the risk reduction tests. The results of one case, accompanied by CFD simulations, are presented in this paper. Experimental results suggest the wall mounted wind tunnel model produces a thicker boundary layer on the fuselage than the CFD predictions, resulting in a larger wing horseshoe vortex suppressing the side of body separation in the juncture flow region. Compared to experimental results, CFD predicts a thinner boundary layer on the fuselage generates a weaker wing horseshoe vortex resulting in a larger side of body separation.

Roozeboom, Nettie H.↗
DOE OSTI2020

Evaporation Sub-model Development for Volume of Fluid (eVOF) Method Applicable to Spray-Wall Interaction Including Film Characteristics with Validation at High Pressure and Temperature Conditions (Final Report)

Internal combustion engines have seen a great evolution over the last several decades through application of high pressure direct injection, multiple injections, and other technologies to reduced fuel consumption, NOx, and PM. Although combustion systems with advanced injection strategies have been studied extensively, there exists a significant fundamental knowledge gap on the fuel-spray interactions with the piston surface and chamber walls. Advanced computational codes validated with experimental techniques have to be developed for accurate representation of the drop impingement, fuel film formation, and vaporization. Current engine CFD spray models utilize a Lagrangian framework for modeling which lacks critical considerations of the physics pertaining to these interactions and thus requiring extensive parameterization, tuning and validation. The team from Michigan Technological University, University of Massachusetts Dartmouth, and Argonne National Laboratory is composed of experts in sprays, combustion, engines and CFD with a wide spectrum of knowledge including specific expertise in the area under consideration. In the proposed work, a VOF modeling approach has been adopted for the spray-wall interaction, film formation and spreading, and vaporization. With the inclusion of a vaporization submodel, a more predictive and accurate simulation of the spray-film was performed without extensive need of parameterization and tuning. Extensive experimentation of the spray-wall interaction under the range of conditions matching the thermodynamic and surface temperatures that occur in diesel and gasoline engines were conducted to validate the SWI submodels and for development of the evaporation sub-model, which has been implemented in the Converge software.

42 ENGINEERING↗
NASA NTRS1974

Effect of forward velocity on the noise characteristics of dual-flow jet nozzles

Acoustic results of jet noise tests conducted on scale-model nozzles at a wall isolation facility are presented and analyzed. The test fixture featured a triple-flow system for the purpose of simulating freestream flow around hot dual-flow jet nozzles. Test nozzles included 1/20th scale models of the exhaust system for the JT8D-9 engine, power plant for the Boeing 727 airplane (bypass ratio 1) and of the JT8D-109 refan engine (bypass ratio 2). It was found that overall jet noise is reduced by the forward velocity in proportion to the fifth and sixth power of the relative primary jet velocity. Spectral and directivity characteristics are presented.

Reed, D. H.↗
NASA NTRS1987

Flutter study of an advanced composite wing with external stores

A flutter test using a scaled model of an advanced composite wing for a Navy attack aircraft has been conducted in the NASA Langley Research Center Transonic Dynamics Tunnel. The model was a wall-mounted half-span wing with a semi-span of 6.63 ft. The wing had an aspect ratio of 5.31, taper ratio of 0.312, and quarter-chord sweep of 25 degrees. The model was supported in a manner that simulated the load path in the carry-through structure of the aircraft and the symmetric boundary condition at the fuselage centerline. The model was capable of carrying external stores from three pylon locations on the wing. Flutter tests were conducted for the wing with and without external stores. No flutter was encountered for the clean wing at test conditions which simulated the scaled airplane operating envelope. Flutter boundaries were obtained for several external store configurations. The flutter boundaries for the fuel tanks were nearly Mach number independent (occurring at constant dynamic pressure). To study the aerodynamic effect of the fuel tank stores, pencil stores (slender cylindrical rods) which had the same mass and pitch and yaw inertia as the fuel tanks were tested on the model. These pencil store configurations exhibited a transonic dip in the flutter dynamic pressure, indicating that the aerodynamic effect of the actual fuel tanks on flutter was significant. Several flutter analyses methods were used in an attempt to predict the flutter phenomenon exhibited during the wind-tunnel test. The analysis gave satisfactory predictions of flutter for the pencil store configurations, but unsatisfactory correlation for the actual fuel tank configurations.

Cole, Stanley R.↗
DOE OSTI2021

Electroweak-like baryogenesis with new chiral matter

We propose a framework where a phase transition associated with a gauge symmetry breaking that occurs (not far) above the electroweak scale sets a stage for baryogenesis similar to the electroweak baryogenesis in the Standard Model. A concrete realization utilizes the breaking of SU(2) R × U(1) X → U(1) Y . New chiral fermions charged under the extended gauge symmetry have nonzero lepton numbers, which makes the B - L symmetry anomalous. The new lepton sector contains a large flavor-dependent CP violation, similar to the Cabibbo-Kobayashi-Maskawa phase, without inducing sizable electric dipole moments of the Standard Model particles. A bubble wall dynamics associated with the first-order phase transition and SU(2) R sphaleron processes generate a lepton asymmetry, which is transferred into a baryon asymmetry via the ordinary electroweak sphaleron process. Unlike the Standard Model electroweak baryogenesis, the new phase transition can be of the strong first order and the new CP violation is not significantly suppressed by Yukawa couplings, so that the observed asymmetry can be produced. The model can be probed by collider searches for new particles and the observation of gravitational waves. One of the new leptons becomes a dark matter candidate. The model can be also embedded into a left-right symmetric theory to solve the strong CP problem.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
NASA NTRS1968

Survey of the Relationship Between Theory and Experiment for Convective Heat Transfer From Rocket Combustion Gases

The problem of predicting the distribution of heat flux in a rocket engine is complicated by a flow field and thermodynamics so perturbed by as yet indescribable combustion effects that it continues to be necessary to utilize a simplified theoretical model against which to compare experimental results in order to gain insight. The most successful model applied to this problem is that of a compressible, turbulent, accelerating boundary layer flow with an arbitrary point of origin, and chemical reaction restricted to equilibrium recombination or diffusion-limited recombination on a catalytic wall. Analysis of this model was not possible without making several assumptions about the nature of the skin friction coefficient relationship and Reynold's analogy in an accelerating boundary layer flow. The paper summarizes the analysis of this model (together with detailed treatment of the special problems of variable properties, chemical reaction, and sources of transport properties), and describes several closed-form equations that have also been developed and applied to the problem. Described are results of several experiments with hot air flow made for the purpose of checking the validity of the assumptions in the boundary layer analysis. Most significant, however, are presentations of sample distributions of local heat flux measured in rocket engines operating with a wide range of conditions, propellants, injectors and sizes. When compared with predictions, these data show the significant range of deviation in heat flux that results from the combustion effects and which precludes a generally accurate prediction theory or correlation equation until the combustion process itself is correlated.

D R Bartz↗
NASA NTRS1989

Control of separated flow past a cylinder using tangential wall jet blowing

A theoretical analysis is conducted to study the effect of a tangential wall jet on the control of two dimensional separated flow past a circular cylinder. For a tangential wall jet, the mathematical model derived previously is used and the vortex cloud method is adopted for the calculation of the external flow field. For certain limiting cases, the governing equations are simplified and closed forms of solutions for the wall jet parameters can be obtained. It is observed that the wall jet is very efficient in reducing drag by delaying the separation point in the range of small blowing strength. The suction force induced by the wall jet is negligible to the drag due to the external stream. However this suction increases the drag when the blowing strength is large.

Oh, Sejong↗
NASA NTRS2006

Accuracy Quantification of the Loci-CHEM Code for Chamber Wall Heat Transfer in a GO2/GH2 Single Element Injector Model Problem

A robust rocket engine combustor design and development process must include tools which can accurately predict the multi-dimensional thermal environments imposed on solid surfaces by the hot combustion products. Currently, empirical methods used in the design process are typically one dimensional and do not adequately account for the heat flux rise rate in the near-injector region of the chamber. Computational Fluid Dynamics holds promise to meet the design tool requirement, but requires accuracy quantification, or validation, before it can be confidently applied in the design process. This effort presents the beginning of such a validation process for the Loci-CHEM CFD code. The model problem examined here is a gaseous oxygen (GO2)/gaseous hydrogen (GH2) shear coaxial single element injector operating at a chamber pressure of 5.42 MPa. The GO2/GH2 propellant combination in this geometry represents one the simplest rocket model problems and is thus foundational to subsequent validation efforts for more complex injectors. Multiple steady state solutions have been produced with Loci-CHEM employing different hybrid grids and two-equation turbulence models. Iterative convergence for each solution is demonstrated via mass conservation, flow variable monitoring at discrete flow field locations as a function of solution iteration and overall residual performance. A baseline hybrid was used and then locally refined to demonstrate grid convergence. Solutions were obtained with three variations of the k-omega turbulence model.

West, Jeff↗
NASA NTRS2006

Accuracy Quantification of the Loci-CHEM Code for Chamber Wall Heat Fluxes in a G02/GH2 Single Element Injector Model Problem

A robust rocket engine combustor design and development process must include tools which can accurately predict the multi-dimensional thermal environments imposed on solid surfaces by the hot combustion products. Currently, empirical methods used in the design process are typically one dimensional and do not adequately account for the heat flux rise rate in the near-injector region of the chamber. Computational Fluid Dynamics holds promise to meet the design tool requirement, but requires accuracy quantification, or validation, before it can be confidently applied in the design process. This effort presents the beginning of such a validation process for the Loci- CHEM CPD code. The model problem examined here is a gaseous oxygen (GO2)/gaseous hydrogen (GH2) shear coaxial single element injector operating at a chamber pressure of 5.42 MPa. The GO2/GH2 propellant combination in this geometry represents one the simplest rocket model problems and is thus foundational to subsequent validation efforts for more complex injectors. Multiple steady state solutions have been produced with Loci-CHEM employing different hybrid grids and two-equation turbulence models. Iterative convergence for each solution is demonstrated via mass conservation, flow variable monitoring at discrete flow field locations as a function of solution iteration and overall residual performance. A baseline hybrid grid was used and then locally refined to demonstrate grid convergence. Solutions were also obtained with three variations of the k-omega turbulence model.

West, Jeff↗