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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 613 records · Page 34
All sourcesNASA NTRSDOE OSTI
DOE OSTI2023

Modeling ash deposition and shedding during oxy-combustion of coal/rice husk blends at 70% inlet O2

Abstract Co-firing rice husk (RH) and coal with carbon capture using oxy-combustion presents a net carbon negative energy production opportunity. In addition, the high fusion temperature of the non-sticky, silica rich, RH can mitigate ash deposition as well as promote shedding of deposits. To identify the optimum operating conditions, fuel particle sizes, and blend ratios that minimize ash deposition, a Computational Fluid Dynamic methodology with add-on ash deposition and shedding models were employed to predict outer ash deposition and shedding rates during co-combustion of coal/RH in AIR and O 2 /CO 2 (70/30 vol%, OXY70) oxidizer compositions. After ensuring that the fly-ash particle size distributions and particle Stokes numbers near the deposition surface were accurately represented (to model impaction), appropriate models for coal ash and RH ash viscosities that were accurate in the temperature region (1200–1300 K) of interest in this study were identified. A particle viscosity and kinetic energy (PKE) based capture criterion was enforced to model the ash capture. An erosion/shedding criterion that takes the deposit melt fraction and the energy consumed during particle impact into account was also implemented. Deposition rate predictions as well as the deposition rate enhancement (OXY70/AIR) were in good agreement with measured values. While the OXY70 scenario was associated with a significant reduction (60%–70%) in flue gas velocities, it also resulted in larger fly-ash particles. As a result, the PKE distributions of the erosive RH ash were similar in both scenarios and resulted in similar shedding rates.

Energy & Fuels↗
NASA NTRS1995

Triggering of longitudinal combustion instabilities in solid rocket motors: Nonlinear combustion response

Pulsed oscillations in solid rocket motors are investigated with emphasis on nonlinear combustion response. The study employs a wave equation governing the unsteady motions in a two-phase flow, and a solution technique based on spatial- and time-averaging. A wide class of combustion response functions is studied to second-order in fluctuation amplitude to determine if, when, and how triggered instabilities arise. Conditions for triggering are derived from analysis of limit cycles, and regions of triggering are found in parametric space. Based on the behavior of model dynamical systems, introduction of linear cross-coupling and quadratic self-coupling among the acoustic modes appears to be the manner in which the nonlinear combustion response produces triggering to a stable limit cycle. Regions of initial conditions corresponding to stable pulses were found, suggesting that stability depends on initial phase angle and harmonic content, as well as the composite amplitude, of the pulse.

Wicker, J. M.↗
DOE OSTI2024

Simulations of flaming combustion and flaming-to-smoldering transition in wildland fire spread at flame scale

Our objective in the present study is to provide basic insights into the coupling between external-gas and solid biomass vegetation processes that control the dynamics of flame spread in wildland fire problems. Here we focus on a modeling approach that resolves processes occurring at vegetation and flame scales, i.e., the formation of flammable vapors due to the thermal degradation of the solid biomass, the subsequent combustion in ambient air, the thermal feedback to the biomass through radiative and convective heat transfer, and the possible transition from flaming combustion (taking place outside of the solid biomass) to smoldering combustion (taking place inside the solid biomass). The capability uses a multiphase combustion framework and treats external-gas processes through a Large Eddy Simulation solver and solid biomass processes through a discrete particle model. The discrete particle model adopts a one-dimensional porous medium formulation, includes descriptions of drying, thermal pyrolysis, oxidative pyrolysis, and char oxidation, as well as a description of the external-gas-to-solid-biomass diffusion of oxygen mass; the discrete particle model thereby provides a treatment of in-depth oxidative processes and allows the simulation of smoldering combustion. The modeling capability is applied to the simulation of fire spread across a surrogate biomass vegetation bed corresponding to a discrete array of cylindrical-shaped, vertically-oriented, pine wood sticks, characterized by a monomodal size distribution, in horizontal flat terrain and under wind-aided conditions. The numerical results demonstrate that the model can simulate successful flaming-to-smoldering transition followed by complete biomass consumption.

01 COAL, LIGNITE, AND PEAT↗
NASA NTRS1993

Microgravity flight research in a dynamic and changing environment

NASA's Microgravity Science and Applications Division conducts both basic and applied low-g research encompassing biotechnology, fluid physics, combustion, and materials science, in industry, university, and government research settings. The present evaluation of such a diverse research program's management gives attention to the fiscal constraints, polical factors, technological changes, international cooperation requirements, and educational responsibilities that must be addressed.

Schmitz, Robert A.↗
NASA NTRS1995

Dynamic response of a pulsed Burke-Schumann diffusion flame

Turbulent flames are often envisioned as an ensemble of random vortices interacting with the combustion process. A better understanding of the vortex-flame interactions therefore would be useful in improving the modeling of turbulent diffusion flames. Substantial simplification may be made by investigating controlled interactions in a laminar flame, as opposed to random interactions in a turbulent flame. The general goals of the research project are to improve our understanding of (1) the influence of buoyancy on co-flow diffusion flames and (2) the effects of buoyancy on vortex-flame interactions in co-flow diffusion flames. As a first step toward objective (2), we conducted a joint experimental and numerical investigation of the vortex-flame interaction. Vortices were produced by mechanically pulsing the fuel flow at a low frequency, e.g., 10 Hz. Experiments were conducted using a nonflickering Burke-Schumann flame in both microgravity (mu-g) and normal gravity (1g) as a means of varying the buoyant force without modification of the pressure (i.e., density). The effects of buoyant convection may then be determined by a comparison of the mu-g and 1g results. The mu-g results may also reveal the important mechanisms which are masked or overwhelmed by buoyant convection in 1g. A numerical investigation was conducted using a validated, time-accurate numerical code to study the underlying physics during the flame interaction and to assist the interpretation of the experimental results.

Sheu, Jyh-Cherng↗
NASA NTRS2018

Performance, Stability and Compatibility of Oxygen/RP-1 Multi-Element Oxidizer-Rich Staged-Combustion Injectors

In 2015 and 2016, the National Aeronautics and Space Administration Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/RP-1 propellant multi-element oxidizer-rich staged-combustion test article. The main objective was to provide thrust chamber combustion stability data as part of the Combustion Stability Tool Development program, although demonstration of performance and compatibility of oxidizer-rich main injectors was also important. Funding was provided by the Air Force Space and Missile Systems Center. Five configurations of main injectors were designed and fabricated, using conventional gas-centered swirl coaxial injector element designs generally similar to those used in oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines. Variations of element features included element size, recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. Ablative combustion chambers were fabricated based on hardware previously used at the NASA MSFC for testing at similar size and pressure. Existing oxygen/RP-1 oxidizer-rich subscale preburner injectors and hot gas ducts from a previous NASA-funded program were modified for use to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Testing of the resulting integrated test article - which included the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was conducted at Test Stand 116 at the East Test Area of the NASA MSFC. The test article was well instrumented with static and dynamic pressure, temperature, and vibration sensors. This paper presents and discusses all the hot-fire test results of the integrated test article thrust chamber. Eighteen successful hot-fire tests of the integrated rig were conducted. Testing was accomplished with all five of the injector element concepts. Main combustion chamber pressures ranged from 710 to 2350 psia, and main combustion chamber mixture ratios ranged from 2.47 to 2.87. A chamber barrier fuel film coolant of about 2% to 4% of the total fuel flow was used for most tests. Characteristic exhaust velocity efficiency excluding the influence of the fuel film cooling ranged from 91% to 98% of theoretical. All tests of the thrust chamber exhibited stable combustion, even down to 40% of nominal operating pressures. Compatibility of the injector face and combustion chamber walls was acceptable. This paper is a follow-on to publication of preliminary test data presented at the 2016 JANNAF Liquid Propulsion Subcommittee meeting.

Hulka, J. R.↗
DOE OSTI2025

Insights into oxygen diffusion in rare earth disilicate environmental barrier coatings

Environmental barrier coatings (EBC) are crucial for the use of SiC-based ceramic matrix composites in high-temperature combustion environments, yet knowledge of oxygen diffusion in these coatings is limited. This study investigates oxygen diffusion dynamics in the β-RE 2 Si 2 O 7 system to minimize oxygen penetration in rare earth disilicates. We analyze defect formation energy under varying oxygen conditions, identifying key diffusion mechanisms. In oxygen-rich environments, the most favorable neutral interstitial oxygen diffuses along the [110] direction. In oxygen-poor conditions, neutral oxygen vacancies rotate around Y and Si atoms, exhibiting a diffusivity of 6.59×10 −22 m 2 /s at 1500 K for β-Yb 2 Si 2 O 7 . Under intermediate oxygen levels, charged interstitial oxygen diffuses via concerted interstitialcy along the [001] direction with a diffusivity of 6.21×10 −17 m 2 /s. Additionally, alloying rare earth Y with Er and Yb increases diffusion barriers, contributing to improved EBC performance in extreme environments. The insights gained provides valuable guidance for designing robust coatings tailored to withstand extreme operational environments.

36 MATERIALS SCIENCE↗
NASA NTRS1972

Characterization of the space shuttle reaction control system engine

A computer program was developed and written in FORTRAN 5 which predicts the transient and steady state performance and heat transfer characteristics of a pulsing GO2/GH2 rocket engine. This program predicts the dynamic flow and ignition characteristics which, when combined in a quasi-steady state manner with the combustion and mixing analysis program, will provide the thrust and specific impulse of the engine as a function of time. The program also predicts the transient and steady state heat transfer characteristics of the engine using various cooling concepts. The computer program, test case, and documentation are presented. The program is applicable to any system capable of utilizing the FORTRAN 4 or FORTRAN 5 language.

Wilson, M. S.↗
NASA NTRS1990

CARS Temperature Measurements in a Hypersonic Propulsion Test Facility

Nonintrusive diagnostic measurements were performed in the supersonic reacting flow of the Hypersonic Propulsion Test Cell 2 at NASA-Langley. A Coherent Anti-stokes Raman Spectroscopy (CARS) system was assembled specifically for the test cell environment. System design considerations were: (1) test cell noise and vibration; (2) contamination from flow field or atmospheric borne dust; (3) unwanted laser or electrically induced combustion (inside or outside the duct); (4) efficient signal collection; (5) signal splitting to span the wide dynamic range present throughout the flow field; (6) movement of the sampling volume in the flow; and (7) modification of the scramjet model duct to permit optical access to the reacting flow with the CARS system. The flow in the duct was a nominal Mach 2 flow with static pressure near one atmosphere. A single perpendicular injector introduced hydrogen into the flow behind a rearward facing step. CARS data was obtained in three planes downstream of the injection region. At least 20 CARS data points were collected at each of the regularly spaced sampling locations in each data plane. Contour plots of scramjet combustor static temperature in a reacting flow region are presented.

Jarrett, Olin, Jr.↗
NASA NTRS1992

A comparative study of the effects of inhibitor stub length on solid rocket motor combustion chamber pressure oscillations: RSRM at T = 80 seconds, preliminary results

Results from a continuing, time-accurate computational study of the combustion gas flow inside the Space Shuttle Redesigned Solid Rocket Motor (RSRM) are presented. These computational fluid dynamic (CFD) analyses duplicate unsteady flow effects which interact in the RSRM to produce pressure oscillations, and resulting thrust oscillations, at nominally 15, 30, and 45 Hz. Results of the Navier-Stokes computations made at mean pressure and flow conditions corresponding to 80 seconds after motor ignition both with and without a protruding, rigid inhibitor at the forward joint cavity are presented here.

Chasman, D.↗
NASA NTRS1992

The numerical dynamic for highly nonlinear partial differential equations

Problems associated with the numerical computation of highly nonlinear equations in computational fluid dynamics are set forth and analyzed in terms of the potential ranges of spurious behaviors. A reaction-convection equation with a nonlinear source term is employed to evaluate the effects related to spatial and temporal discretizations. The discretization of the source term is described according to several methods, and the various techniques are shown to have a significant effect on the stability of the spurious solutions. Traditional linearized stability analyses cannot provide the level of confidence required for accurate fluid dynamics computations, and the incorporation of nonlinear analysis is proposed. Nonlinear analysis based on nonlinear dynamical systems complements the conventional linear approach and is valuable in the analysis of hypersonic aerodynamics and combustion phenomena.

Lafon, A.↗
NASA NTRS1993

Development of a droplet breakup model considering aerodynamic and droplet collision effects

A model is currently under development to predict the occurrence and outcome of spray droplet breakup induced by aerodynamic forces and droplet collisions. It is speculated that these phenomena may be significant in determining the droplet size distribution in a spray subjected to acoustic velocity fluctuations. The goal is to integrate this breakup model into a larger spray model in order to examine the effects of combustion instabilities on liquid rocket motor fuel sprays. The model is composed of three fundamental components: a dynamic equation governing the deformation of the droplet, a criterion for breakage based on the amount of deformation energy stored in the droplet and an energy balance based equation to predict the Sauter mean diameter of the fragments resulting from breakup. Comparison with published data for aerodynamic breakup indicates good agreement in terms of predicting the occurrence of breakup. However, the model significantly over predicts the size of the resulting fragments. This portion of the model is still under development.

Wert, K. L.↗
NASA NTRS1995

Computational fluid dynamic analysis of hybrid rocket combustor flowfields

Computational fluid dynamic analyses of the Navier-Stokes equations coupled with solid-phase pyrolysis, gas-phase combustion, turbulence and radiation are performed to study hybrid rocket combustor flowfields. The computational study is closely co-ordinated with a companion experimental program using a planar slab burner configuration with HTPB as fuel and gaseous oxygen. Computational predictions agree reasonably well with measurement data of fuel regression rates and surface temperatures. Additionally, most of the parametric trends predicted by the model are in general agreement with experimental trends. The computational model is applied to extend the results from the lab-scale to a full-scale axisymmetric configuration. The numerical predictions indicate that the full-scale configuration burns at a slower rate than the lab-scale combustor under identical specific flow rate conditions. The results demonstrate that detailed CFD analyses can play a useful role in the design of hybrid combustors.

Venkateswaran, S.↗
NASA NTRS2001

Acoustical and Flowfield Characterization of a Scaled Tabletop Rocket

An analysis of the acoustical and flowfield environment for the scaled 1-pound-force (lbf) thrust tabletop motor was performed. This tabletop motor from NASA Stennis Space Center Is composed of Plexiglas burning In gaseous oxygen with a graphite insert for the nozzle portion. The nozzle has a throat diameter of 0.2 inch and an exit diameter of 0.38 Inch. With a chamber pressure at 55 pounds per square Inch absolute (psia), a normal shock is formed immediately downstream of the nozzle exit plane as the combustion products exhaust into the ambient at atmospheric pressure. The jet characterization Is based on computational fluid dynamics (CFD) in conjunction with Kirchhoff surface integral formulation and compared with correlations developed for measured rocket noise and a pressure fluctuation scaling (PFS) method. Predictions and comparisons are made for the overall sound pressure levels (OASPL's) and spectral dependence of sound pressure level (SPL). The overall objective of this effort is to develop methods for scaling the acoustic and flowfield environment of rockets with a wide range of thrust (1 lbf to 1 million lbf).

Kandula, Max↗
NASA NTRS2016

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Jones, Daniel R.↗
DOE OSTI2022

Mechanistic study of chemical looping reactions between solid carbon fuels and CuO

Copper (Cu) based chemical-looping combustion (CLC) is a promising process that utilizes solid carbon fuel such as coal and biomass. Understanding the reaction kinetics in this process can facilitate the industrial design of CLC units. Here, in this work, molecular dynamics (MD) simulations were performed to investigate the reaction kinetics of n-butane and two simplified solid carbon fuels (lignite and anthracite) with/without copper oxide (CuO) nanoparticle. In addition, experiments were conducted on the thermal characteristics, flammability, and flame speeds of CuO and solid carbon mixtures. For n-butane, oxidation on the CuO surface is the primary reaction since the activation energy of the surface reaction is much lower than that of oxygen (O 2 ) combustion (9.2 vs. 53.3 kcal/mol). On the other hand, for lignite, there is a smaller difference in the activation energy of O 2 combustion and surface reaction (23.04 vs. 6.34 kcal/mol). We hypothesize that some solid carbon fuels can have different reaction kinetics dependent upon temperature. This is proven by the case of a simplified anthracite coal: an Arrhenius plot shows that this solid carbon fuel has two different reaction kinetic regimes and the critical temperature for the change in kinetics is related to the oxygen uncoupling of the CuO nanoparticle. Like the modeling simulations, a change in activation energy is observed in the experimental results, where desorption of molecular oxygen from CuO becomes important.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1990

Comparisons of a Three-Dimensional, Full Navier Stokes Computer Model with High Mach Number Combuster Test Data

Comparisons between scramjet combustor data and a three-dimensional full Navier-Stokes calculation have been made to verify and substantiate computational fluid dynamics (CFD) codes and application procedures. High Mach number scramjet combustor development will rely heavily on CFD applications to provide wind tunnel-equivalent data of quality sufficient to design, build and fly hypersonic aircraft. Therefore. detailed comparisons between CFD results and test data are imperative. An experimental case is presented, for which combustor wall static pressures were measured and flow-fieid interferograms were obtained. A computer model was done of the experiment, and counterpart parameters are compared with experiment. The experiment involved a subscale combustor designed and fabricated for the National Aero-Space Plane Program, and tested in the Calspan Corporation 96" hypersonic shock tunnel. The combustor inlet ramp was inclined at a 20 angle to the shock tunnel nozzle axis, and resulting combustor entrance flow conditions simulated freestream M=10. The combustor body and cowl walls were instrumented with static pressure transducers, and the combustor lateral walls contained windows through which flowfield holographic interferograms were obtained. The CFD calculation involved a three-dimensional time-averaged full Navier-Stokes code applied to the axial flow segment containing fuel injection and combustion. The full Navier-Stokes approach allowed for mixed supersonic and subsonic flow, downstream-upstream communication in subsonic flow regions, and effects of adverse pressure gradients. The code included hydrogen-air chemistry in the combustor segment which begins near fuel injection and continues through combustor exhaust. Combustor ramp and inlet segments on the combustor lateral centerline were modelled as two dimensional. Comparisons to be shown include calculated versus measured wall static pressures as functions of axial flow coordinate, and calculated path-averaged density contours versus an holographic Interferogram.

Watkins, William B.↗
DOE OSTI2021

Determination of a most representative cycle from cylinder pressure ensembles via statistical method using distribution skewness

In internal combustion engine research, cylinder pressure measurements provide valuable information about the underlying thermodynamic and combustion processes, and are typically collected in ensembles of several 100 traces. Although in some particular fields of combustion research all traces are analyzed, in most cases only one trace is studied because analyzing all the traces is impractical due to the large number of collected samples. Instead, an ensemble-averaged pressure trace is commonly calculated and used for analysis. However, this pressure trace is highly smoothed and dynamic information is lost during the averaging process. With the average trace, pressure rise rates are lower and pressure oscillations such as the ones resulting from combustion knock are lost. In this work, a statistical method was developed to determine the “most representative cycle,” which is the cycle from the ensemble that has the pressure trace most representative of the engine operating condition. Eleven characteristic parameters are computed from each pressure trace and probabilistic distributions are obtained for each of the parameters using all the traces in the ensemble. Finally, the most representative cycle is selected by means of a cost function minimization. The benefits of this method are illustrated using experimental data from four very different engine platforms, under four different combustion modes and over a range of operating conditions.

42 ENGINEERING↗