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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Numerical Simulation of a High Mach Number Jet Flow

The recent efforts to develop accurate numerical schemes for transition and turbulent flows are motivated, among other factors, by the need for accurate prediction of flow noise. The success of developing high speed civil transport plane (HSCT) is contingent upon our understanding and suppression of the jet exhaust noise. The radiated sound can be directly obtained by solving the full (time-dependent) compressible Navier-Stokes equations. However, this requires computational storage that is beyond currently available machines. This difficulty can be overcome by limiting the solution domain to the near field where the jet is nonlinear and then use acoustic analogy (e.g., Lighthill) to relate the far-field noise to the near-field sources. The later requires obtaining the time-dependent flow field. The other difficulty in aeroacoustics computations is that at high Reynolds numbers the turbulent flow has a large range of scales. Direct numerical simulations (DNS) cannot obtain all the scales of motion at high Reynolds number of technological interest. However, it is believed that the large scale structure is more efficient than the small-scale structure in radiating noise. Thus, one can model the small scales and calculate the acoustically active scales. The large scale structure in the noise-producing initial region of the jet can be viewed as a wavelike nature, the net radiated sound is the net cancellation after integration over space. As such, aeroacoustics computations are highly sensitive to errors in computing the sound sources. It is therefore essential to use a high-order numerical scheme to predict the flow field. The present paper presents the first step in a ongoing effort to predict jet noise. The emphasis here is in accurate prediction of the unsteady flow field. We solve the full time-dependent Navier-Stokes equations by a high order finite difference method. Time accurate spatial simulations of both plane and axisymmetric jet are presented. Jet Mach numbers of 1.5 and 2.1 are considered. Reynolds number in the simulations was about a million. Our numerical model is based on the 2-4 scheme by Gottlieb & Turkel. Bayliss et al. applied the 2-4 scheme in boundary layer computations. This scheme was also used by Ragab and Sheen to study the nonlinear development of supersonic instability waves in a mixing layer. In this study, we present two dimensional direct simulation results for both plane and axisymmetric jets. These results are compared with linear theory predictions. These computations were made for near nozzle exit region and velocity in spanwise/azimuthal direction was assumed to be zero.

Hayder, M. Ehtesham↗
NASA NTRS1993

A numerical investigation of a subsonic jet in a crossflow

The flowfield induced by a single, subsonic jet exhausting perpendicularly from a flat plate into a subsonic crossflow has been numerically investigated. The test case was chosen to match available experimental data where the jet Mach number was 0.78, and the freestream Mach number was 0.13. Time-averaged solutions were obtained using the thin-layer Navier-Stokes equations and two overlapping grids. The solutions were sensitive to the radial grid clustering near the edge of the jet and to the far-field boundary conditions. Experimental data comparisons were required to determine the most appropriate jet grid and satisfactory boundary conditions. Globally, the solutions converged in about 6000 iterations. The computational results accurately showed the deflected jet and associated contrarotating vortices. The fine clustered grid in the region upstream of the jet exit allowed the horseshoe vortex in the boundary layer near the jet exit to be captured. Most importantly for aircraft applications, the computed plate pressure distributions compared favorably with the experimental data over most of the surface. However, in the wake region immediately downstream of the jet exit, where there is extensive flow separation, some discrepancies with experimental data were observed. Two turbulence models were used in this study: (1) the zero-equation, two layer Baldwin-Lomax turbulence model; and (2) one-equation Baldwin-Barth turbulence model. The turbulence models gave results which generally compared no better with experimental data than the laminar computation results.

Chiu, Stephen H.↗
DOE OSTI2021

Progress of Divertor Heat and Particle Flux Control in EAST for Advanced Steady-State Operation in the Last 10 Years

Active control of the excessively high heat and particle fluxes on the divertor target plates is of fundamental importance to the steady state operation of tokamaks, especially for fusion reactors. A series of experiments have been carried out on this critical issue to relieve the plasma-wall interactions in the experimental advanced superconducting tokamak (EAST) in the last ten years, not only contributing to the long pulse operation of EAST itself, but also providing physical understandings and potential techniques to the next-generation devices like ITER. Here we have characterized the power deposition pattern and broadened the divertor footprint width effectively. The plasma-wetted area is actively handled using either 3-dimentional edge magnetic topology or advanced plasma equilibrium, thereby peak heat flux around the strike point is reduced. Active control of detachment or radiation compatible with core plasma performance has progressed significantly in very recent years, with a series of active feedback control modules developed and utilized successfully, based on the divertor physics advances with both experiments and simulation. The upper divertor of EAST was upgraded from graphite to active water-cooling ITER-like tungsten in 2014, exhibiting much enhanced heat removal capability. As for the particle exhaust including both fueling and impurity particles, in addition to wall conditioning and impurity source control, the efficiency of particle flux exhaust is optimized by making full use of the divertor closure and the plasma drifts in both scrape-off layer and divertor volume. These heat and particle exhaust advances contribute greatly to a series of EAST achievements like H-mode operation over 100 s. A brief near-term plan on the integrated control of divertor plasma-wall interactions in long-time scale will also be introduced, aiming to provide favorable divertor operation solution for ITER and CFETR.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1975

Duct acoustics and acoustic finite element method

A comprehensive solution method - Acoustic Finite Element Method - has been proposed. In order to generate confidence in the method proposed, the problem of influence of shear flow in a two-dimensional channel on sound propagation from an arbitrary source distribution has been theoretically formulated. From the preliminary investigation of application of acoustic finite element method, it seems that it can provide solutions to problems concerning the transmission of sound in variable area ducts, or in ducts in which the boundary layer depends on the streamwise variable, or in ducts with discontinuities in lining impedance and with end termination impedances, both the inlet and exhaust jet flows being taken into account.

Kapur, A.↗
NASA NTRS1995

An approximate theoretical method for modeling the static thrust performance of non-axisymmetric two-dimensional convergent-divergent nozzles

An analytical/numerical method has been developed to predict the static thrust performance of non-axisymmetric, two-dimensional convergent-divergent exhaust nozzles. Thermodynamic nozzle performance effects due to over- and underexpansion are modeled using one-dimensional compressible flow theory. Boundary layer development and skin friction losses are calculated using an approximate integral momentum method based on the classic karman-Polhausen solution. Angularity effects are included with these two models in a computational Nozzle Performance Analysis Code, NPAC. In four different case studies, results from NPAC are compared to experimental data obtained from subscale nozzle testing to demonstrate the capabilities and limitations of the NPAC method. In several cases, the NPAC prediction matched experimental gross thrust efficiency data to within 0.1 percent at a design NPR, and to within 0.5 percent at off-design conditions.

Hunter, Craig A.↗
NASA NTRS1989

Comparison of 3D computation and experiment for non-axisymmetric nozzles

Three dimensional solutions of a single expansion ramp nozzle are computed with the existing PARC computer code by solving the full Navier-Stokes equations. The computations are performed to simulate the non-axisymmetric nozzle flowfield in both the internal/external expansion regions and the exhaust plume in a quiescent ambient environment. Two different configurations of the nozzle at a pressure ratio NPR = 10 are examined. Numerical results of laminar flows are presented, and the wall pressure distributions are compared with the experimental data.

Lai, H.↗
NASA NTRS1989

Comparison of 3D computation and experiment for non-axisymmetric nozzles

Three-dimensional solutions of a single expansion ramp nozzle are computed with the existing PARC computer code by solving the full Navier-Stokes equations. The computations are performed to simulate the non-axisymmetric nozzle flowfield in both the internal/external expansion regions and the exhaust plume in a quiescent ambient environment. Two different configurations of the nozzle at a pressure ratio NPR = 10 are examined. Numerical results of laminar flows are presented, and the wall pressure distributions are compared with the experimental data.

Lai, H.↗
NASA NTRS2005

First Test of Fan Active Noise Control (ANC) Completed

With the advent of ultrahigh-bypass engines, the space available for passive acoustic treatment is becoming more limited, whereas noise regulations are becoming more stringent. Active noise control (ANC) holds promise as a solution to this problem. It uses secondary (added) noise sources to reduce or eliminate the offending noise radiation. The first active noise control test on the low-speed fan test bed was a General Electric Company system designed to control either the exhaust or inlet fan tone. This system consists of a "ring source," an induct array of error microphones, and a control computer. Fan tone noise propagates in a duct in the form of spinning waves. These waves are detected by the microphone array, and the computer identifies their spinning structure. The computer then controls the "ring source" to generate waves that have the same spinning structure and amplitude, but 180 out of phase with the fan noise. This computer generated tone cancels the fan tone before it radiates from the duct and is heard in the far field. The "ring source" used in these tests is a cylindrical array of 16 flat-plate acoustic radiators that are driven by thin piezoceramic sheets bonded to their back surfaces. The resulting source can produce spinning waves up to mode 7 at levels high enough to cancel the fan tone. The control software is flexible enough to work on spinning mode orders from -6 to 6. In this test, the fan was configured to produce a tone of order 6. The complete modal (spinning and radial) structure of the tones was measured with two builtin sets of rotating microphone rakes. These rakes provide a measurement of the system performance independent from the control system error microphones. In addition, the far-field noise was measured with a semicircular array of 28 microphones. This test represents the first in a series of tests that demonstrate different active noise control concepts, each on a progressively more complicated modal structure. The tests are in preparation for a demonstration on a flight-type engine.

Source record↗
NASA NTRS1976

Non-equilibrium radiation from viscous chemically reacting two-phase exhaust plumes

A knowledge of the structure of the rocket exhaust plumes is necessary to solve problems involving plume signatures, base heating, plume/surface interactions, etc. An algorithm is presented which treats the viscous flow of multiphase chemically reacting fluids in a two-dimensional or axisymmetric supersonic flow field. The gas-particle flow solution is fully coupled with the chemical kinetics calculated using an implicit scheme to calculate chemical production rates. Viscous effects include chemical species diffusion with the viscosity coefficient calculated using a two-equation turbulent kinetic energy model.

Penny, M. M.↗
NASA NTRS1993

Numerical simulation of the Space Shuttle Launch Vehicle flowfield with real gas solid rocket plume effects

A numerical simulation of the external ascent flowfield of the Space Shuttle Launch Vehicle (SSLV) was carried out at the freestream Mach number 1.25, the angle of attack -5.1 deg, and the flight Reynolds number 3.25 x 10 exp 6/ft. The simulation is an extension of the solution by Kandula et al. (1991) and incorporates variable gamma effects with a high fidelity grid appropriate for a flight Reynolds number flow solution. Three-dimensional SSLV steady-state simulations with either perfect-gas or variable-gamma-gas Redesigned-Solid-Rocket-Motor (RSRM) plumes are computed on a 14-grid overlapping body-conforming grid system, and the influence of the RSRM exhaust plumes on the SSLV component pressure distributions and integrated loads is examined.

Slotnick, J. P.↗
NASA NTRS1981

Extension to an analysis of turbulent swirling compressible flow for application to axisymmetric small gas turbine ducts

An existing computer program, the Axisymmetric Diffuser Duct Code (ADD code), which calculates compressible turbulent swirling flow through axisymmetric ducts was modified to permit calculation of flows through small gas turbine ducts with struts, guide vanes and large degrees of turning. The improvements include a coordinate generator, an end-wall loss model, and a generalized geometry capability to describe struts and guide vanes in ducts which turn more than 90 degrees. An improved output format was developed to provide the solution on any arbitrary plane in the duct and an extensive literature survey of calculation procedures used in gas turbine technology was completed which suggests improvements in the computer code. Calculations are presented for the flow through the AGT101 small gas turbine inlet duct and turbine exhaust diffuser which demonstrate the ADD code modifications implemented in the investigation. The computed results compare favorably with experimental results.

Anderson, O. L.↗
NASA NTRS1993

Planning collision free paths for two cooperating robots using a divide-and-conquer C-space traversal heuristic

A method was developed to plan feasible and obstacle-avoiding paths for two spatial robots working cooperatively in a known static environment. Cooperating spatial robots as referred to herein are robots which work in 6D task space while simultaneously grasping and manipulating a common, rigid payload. The approach is configuration space (c-space) based and performs selective rather than exhaustive c-space mapping. No expensive precomputations are required. A novel, divide-and-conquer type of heuristic is used to guide the selective mapping process. The heuristic does not involve any robot, environment, or task specific assumptions. A technique was also developed which enables solution of the cooperating redundant robot path planning problem without requiring the use of inverse kinematics for a redundant robot. The path planning strategy involves first attempting to traverse along the configuration space vector from the start point towards the goal point. If an unsafe region is encountered, an intermediate via point is identified by conducting a systematic search in the hyperplane orthogonal to and bisecting the unsafe region of the vector. This process is repeatedly applied until a solution to the global path planning problem is obtained. The basic concept behind this strategy is that better local decisions at the beginning of the trouble region may be made if a possible way around the 'center' of the trouble region is known. Thus, rather than attempting paths which look promising locally (at the beginning of a trouble region) but which may not yield overall results, the heuristic attempts local strategies that appear promising for circumventing the unsafe region.

Weaver, Johnathan M.↗
DOE OSTI2024

Testing of Materials and Coatings at up to 1150°C and 300 bar for use in Oxy-Combustion Turbine

Various materials that may be used in a supercritical carbon dioxide (sCO2) oxy-fuel turbine in the 150-300 MWe size range are being evaluated in various environments of interest. The inlet of the turbine must be capable of 1,150 °C at 300 bar and the exhaust temperature is expected to be within the 725-775 °C range. The design requirements are pushing the known limits of high temperature materials. Little is known about the oxidation of the metallic alloys considered for turbine nozzles and blades or about the resilience of thermal management solutions, such as thermal barrier coatings (TBC), required to accommodate these high temperatures. The combinations of high temperature resistant materials and coatings must be tested in sCO2 to evaluate their reliability in an oxy-fuel turbine environment. A total of 13 alloys were chosen to be tested. They included bare samples, with nano deposited MCrAlY bond coat only, or with the addition of a TBC. A unique test facility was developed to expose those materials in sCO2 at up to 300 bar and 1150 °C for up to 5,000 hours. This was accomplished by placing an induction heater inside an autoclave to achieve those high temperature locally while the pressurized vessel is cooled externally. The specimens are weighed before and after exposure to determine the oxidation rate. The integrity, morphology, and composition of some of the coatings and thermally grown oxide will be investigated by scanning electron microscopy (SEM). This paper presents the up-to-date results of the testing coated and uncoated superalloys in sCO2 at up to 1150 °C and up to 300 bar performed at Southwest Research Institute.

Bocher, Florent↗
NASA NTRS2004

A Model for Shear Layer Effects on Engine Noise Radiation

Prediction of aircraft engine noise is an important aspect of addressing the issues of community noise and cabin noise control. The development of physics based methodologies for performing such predictions has been a focus of Computational Aeroacoustics (CAA). A recent example of code development in this area is the ducted fan noise propagation and radiation code CDUCT-LaRC. Included within the code is a duct radiation model that is based on the solution of FfowcsWilliams-Hawkings (FW-H) equation with a penetrable data surface. Testing of this equation for many acoustic problems has shown it to provide generally better results than the Kirchhoff formula for moving surfaces. Currently, the data surface is taken to be the inlet or exhaust plane for inlet or aft-fan cases, respectively. While this provides reasonable results in many situations, these choices of data surface location lead to a few limitations. For example, the shear layer between the bypass ow and external stream can refract the sound waves radiated to the far field. Radiation results can be improved by including this effect, as well as the rejection of the sound in the bypass region from the solid surface external to the bypass duct surrounding the core ow. This work describes the implementation, and possible approximation, of a shear layer boundary condition within CDUCT-LaRC. An example application also illustrates the improvements that this extension offers for predicting noise radiation from complex inlet and bypass duct geometries, thereby providing a means to evaluate external treatments in the vicinity of the bypass duct exhaust plane.

Nark, Douglas M.↗
NASA NTRS2005

Numerical Prediction of Chevron Nozzle Noise Reduction using Wind-MGBK Methodology

Numerical predictions for single-stream chevron nozzle flow performance and farfield noise production are presented. Reynolds Averaged Navier Stokes (RANS) solutions, produced via the WIND flow solver, are provided as input to the MGBK code for prediction of farfield noise distributions. This methodology is applied to a set of sensitivity cases involving varying degrees of chevron inward bend angle relative to the core flow, for both cold and hot exhaust conditions. The sensitivity study results illustrate the effect of increased chevron bend angle and exhaust temperature on enhancement of fine-scale mixing, initiation of core breakdown, nozzle performance, and noise reduction. Direct comparisons with experimental data, including stagnation pressure and temperature rake data, PIV turbulent kinetic energy fields, and 90 degree observer farfield microphone data are provided. Although some deficiencies in the numerical predictions are evident, the correct farfield noise spectra trends are captured by the WIND-MGBK method, including the noise reduction benefit of chevrons. Implications of these results to future chevron design efforts are addressed.

Engblom, W.A.↗
NASA NTRS2011

Establishing Approaches to Modeling the Ares I-X and Ares I Roll Control System with Free-stream Interaction

Approaches were established for modeling the roll control system and analyzing the jet interactions of the activated roll control system on Ares-type configurations using the USM3D Navier-Stokes solver. Components of the modeling approach for the roll control system include a choice of turbulence models, basis for computing a dynamic equivalence of the real gas rocket exhaust flow in terms of an ideal gas, and techniques to evaluate roll control system performance for wind tunnel and flight conditions. A simplified Ares I-X configuration was used during the development phase of the roll control system modeling approach. A limited set of Navier-Stokes solutions was obtained for the purposes of this investigation and highlights of the results are included in this paper. The USM3D solutions were compared to equivalent solutions at select flow conditions from a real gas Navier- Stokes solver (Loci-CHEM) and a structured overset grid Navier-Stokes solver (OVERFLOW).

Pao, S. Paul↗
DOE OSTI2023

Modeling the U.S. Western Electric Interconnection to Understand the Consequences of Hydrometeorological Extremes and Options for Risk Mitigation

Electricity grid operators around the world face a dual challenge; withstanding increasingly severe weather and the longer term impacts of climate change, while simultaneously decarbonizing. Extreme weather events such as heat waves and droughts are rising in both severity and frequency, which is threatening the reliability of electricity grids through increased demand, generation capacity losses, and equipment failures. Consequently, incorporating hydrometeorological stressors into computational power systems analysis is becoming an even more critical tool in long term planning and short term operations. However, there is a general lack of open-source customizable grid simulation software capable of exhaustively stress testing the grid under hydrometeorological uncertainty, and/or examining potential risk mitigation pathways. A related, persistent challenge for power system modelers is striking an appropriate balance between model fidelity (e.g. spatial scale and time resolution) and computational tractability (wall clock run-time). In this study, we are proposing a solution to this problem with open-source software that allows users to seamlessly customize the scale and track the accuracy of grid operations models. Our approach allows users to search over numerous model parameters (network topology, mathematical formulation, economic hurdle rates, and transmission line scaling) to identify model instantiations that accommodate experimental design. Further, we use this approach to demonstrate the importance of including extreme weather events in model validation and model selection. Focusing on the occurrence of heatwaves and droughts in the U.S. Western Interconnection, we examine role of extreme events in balancing tradeoffs between model fidelity and run-time at the model design stage.

Economics↗
NASA NTRS1956

Internal performance characteristics of short convergent-divergent exhaust nozzles designed by the method of characteristics

Internal performance data on a short exhaust nozzle designed by the method of characteristics were obtained over a range of pressure ratios from 1.5 to 22. The peak thrust coefficient was not affected by a shortened divergent section, but it occurred at lower pressure ratios due to reduction in expansion ratio. This nozzle contour based on characteristics solution gave higher thrust coefficients than a conical convergent-divergent nozzle of equivalent length. Abrupt-inlet sections permitted a reduction in nozzle length without a thrust-coefficient reduction.

Krull, H George↗