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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 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Physics and Modeling of Quasi-Steady and Forced Shock Trains

The dual-mode scramjet is a propulsive concept for contemporary hypersonic vehicles promising flight cruise speeds upwards of Mach 5. Despite decades of development and a recent resurgence in interest, such engines are not routinely deployed for commercial or military purposes. Instead, they have been confined to the realm of ground-test research and limited one-off flight test missions. One of the key stumbling blocks on the critical path to deployment of a dual-mode scramjet is the optimal design and stable operation of the isolator, a typically constant area duct that separates the scramjet inlet from the combustor. During ramjet mode at Mach numbers from 3 to 5 and during the transition to scramjet mode at higher Mach numbers, the isolator contains the pseudo-shock. Necessary to meet the back-pressure in the combustor, the pseudo-shock is typically divided into a shock dominated shock train region followed by a turbulence driven mixing region. Fluctuations in combustor pressure due to vehicle maneuvering, throttling, or combustion instabilities drive the shock train to shift streamwise within the isolator. In response to sharp pressure fluctuations, the shock train is unable to stabilize within the isolator, and the scramjet unstarts by ejecting the shock train out the inlet. This work addresses three key gaps in understanding of the isolator pseudo-shock. The phenomenon is studied in three experimental facilities to investigate the quasi-steady and unsteady fluid dynamics of the shock train in fueled and un-fueled test conditions. The strong link between boundary layer morphology and shock train structure is studied in detail and used as a tool to evaluate the state-of-the-art definitions and models of the pseudo-shock. Identifying the limitations of these models, a new process driven flux-conserved pseudo-shock model is proposed. This provides a computationally efficient model grounded in the fluid dynamic mechanisms of the pseudo-shock with better predictive capability than the current state-of-the-art across all facilities tested. The unsteady behavior of the shock train is further investigated to identify and understand two distinct forced dynamic modes: mechanically induced, low-rate forcing and high-rate combustion driven unsteadiness that sheds upstream propagating shocks. The proposed process driven model is found to be effective in the the limiting case of high-rate combustion driven unsteadiness by evaluating the model in a shock stationary reference frame. Further analysis identifies the approximately rigid chain motion of the shock train structure to mechanically induced forcing. The time delay between the onset of forced dynamics is found to saturate to a constant value and maximum shock train speed along the isolator with respect to the perturbation rate of change. The result of this work provides dual-mode scramjet designers with improved tools and physical understanding to develop more efficient, robust, and capable propulsion systems for hypersonic vehicles.

JAXA↗
NASA NTRS1983

Importance of inlet boundary conditions for numerical simulation of combustor flows

Fluid dynamic computer codes for the mathematical simulation of problems in gas turbine engine combustion systems are required as design and diagnostic tools. To eventually achieve a performance standard with these codes of more than qualitative accuracy it is desirable to use benchmark experiments for validation studies. Typical of the fluid dynamic computer codes being developed for combustor simulations is the TEACH (Teaching Elliptic Axisymmetric Characteristics Heuristically) solution procedure. It is difficult to find suitable experiments which satisfy the present definition of benchmark quality. For the majority of the available experiments there is a lack of information concerning the boundary conditions. A standard TEACH-type numerical technique is applied to a number of test-case experiments. It is found that numerical simulations of gas turbine combustor-relevant flows can be sensitive to the plane at which the calculations start and the spatial distributions of inlet quantities for swirling flows.

Sturgess, G. J.↗
NASA NTRS1984

Measurement of spray combustion processes

A free jet configuration was chosen for measuring noncombusting spray fields and hydrocarbon-air spray flames in an effort to develop computational models of the dynamic interaction between droplets and the gas phase and to verify and refine numerical models of the entire spray combustion process. The development of a spray combustion facility is described including techniques for laser measurements in spray combustion environments and methods for data acquisition, processing, displaying, and interpretation.

Peters, C. E.↗
NASA NTRS1956

Ram-jet Performance

The ram jet is basically one of the most dimple types of aircraft engine. It consists only of an inlet diffuser, a combustion system, and an exit nozzle. A typical ram-jet configuration is shown in figure 128. The engine operates on the Brayton cycle, and ideal cycle efficiency depends only on the ratio of engine to ambient pressure. The increased, engine pressures are obtained by ram action alone, and for this reason the ram jet has zero thrust at zero speed. Therefore, ram-jet-powered aircraft must be boosted to flight speeds close to a Mach number of 1.0 before appreciable thrust is generated by the engine. Since pressure increases are obtained by ram action alone, combustor-inlet pressures and temperatures are controlled by the flight speed, the ambient atmospheric condition, and by the efficiency of the inlet diffuser. These pressures and temperatures, as functions of flight speed and altitude, are shown in figure 129 for the NACA standard atmosphere and for practical values of diffuser efficiency. It can be seen that very wide ranges of combustor-inlet temperatures and pressures may be encountered over the ranges of flight velocity and altitude at which ram jets may be operated. Combustor-inlet temperatures from 500 degrees to 1500 degrees R and inlet pressures from 5 to 100 pounds per square inch absolute represent the approximate ranges of interest in current combustor development work. Since the ram jet has no moving parts in the combustor outlet, higher exhaust-gas temperatures than those used in current turbojets are permissible. Therefore, fuel-air ratios equivalent to maximum rates of air specific impulse or heat release can be used, and, for hydrocarbon fuels, this weight ratio is about 0.070. Lower fuel-air ratios down to about 0.015 may also be required to permit efficient cruise operation. This fuel-air-ratio range of 0.015 to 0.070 used in ram jets can be compared with the fuel-air ratios up to 0.025 encountered in current turbojets. Ram-jet combustor-inlet velocities range from 150 to 400 feet per second. These high linear velocities combined with the relatively low pressure ratios obtainable in ram jets require that the pressure drop through the combustor be kept low to avoid excessive losses in cycle efficiency. It has been estimated that, for a long-range ram-jet engine, an increase in pressure loss of one dynamic head would require a compensating 1-percent increase in combustion efficiency. Therefore, combustor pressure-loss coefficients (pressure drop/impact pressure) of the order of 1 to 4 are found in most current engines. The operating conditions described impose major problems in the design of stable and efficient ram-jet combustion systems. This chapter presents a survey of ram-jet combustor research and, where possible, points out criteria that may be useful in the design of ram-jet combustion systems.

Cervenko, A. J.↗
DOE OSTI2022

Differentiation of static and dynamic interfacial area in the structured packed column

Effective mass-transfer area plays a key role for post-combustion carbon capture. The static area, an inactive part relative to mass transfer, needs to be differentiated in the interfacial area. In the present study, CFD method is employed to investigate static and dynamic interfacial areas in Mellapak 500.Y structured packing. Here, a wide range of physical properties and loading conditions are included to understand their effects on both areas. The result shows that the relationship between interfacial area and interfacial velocity at the local scale can be described by gamma distribution with good accuracy. The influences of various physical properties derived from simulation are compared with existing experiment-based correlations for both static and interfacial areas. The result shows the dominant influence of viscosity on the static area, and the most importance of the influence of contact angle on the interfacial area. This study also highlights the importance of viscosity, which is usually ignored.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1986

Combustion research in the Internal Fluid Mechanics Division

The goal of this research is to bring computational fluid dynamics to a state of practical application for the aircraft engine industry. The approach is to have a strongly integrated computational and experimental program for all the disciplines associated with the gas turbine and other aeropropulsion systems by advancing the understanding of flow physics, heat transfer, and combustion processes. The computational and experimental research is integrated in the following way: the experiments that are performed provide an empirical data set so that physical models can be formulated to describe the processes that are occurring - for example, turbulence or chemical reaction. These experiments also form a data base for those who are doing code development by providing experimental data against which the codes can be verified and assesed. Models are generated as closure to some of the numerical codes, and they also provide physical insight for experiments. At the same time, codes which solve the complete Navier-Stokes equations can be used as a kind of numerical experiment from which far more extensive data can be obtained than ever could be obtained experimentally. This could provide physical insight into the complex processes that are taking place. These codes are also exercised against experimental data to assess the accuracy and applicability of models.

Mularz, Edward J.↗
NASA NTRS1999

FDNS CFD Code Benchmark for RBCC Ejector Mode Operation

Computational Fluid Dynamics (CFD) analysis results are compared with benchmark quality test data from the Propulsion Engineering Research Center's (PERC) Rocket Based Combined Cycle (RBCC) experiments to verify fluid dynamic code and application procedures. RBCC engine flowpath development will rely on CFD applications to capture the multi-dimensional fluid dynamic interactions and to quantify their effect on the RBCC system performance. Therefore, the accuracy of these CFD codes must be determined through detailed comparisons with test data. The PERC experiments build upon the well-known 1968 rocket-ejector experiments of Odegaard and Stroup by employing advanced optical and laser based diagnostics to evaluate mixing and secondary combustion. The Finite Difference Navier Stokes (FDNS) code was used to model the fluid dynamics of the PERC RBCC ejector mode configuration. Analyses were performed for both Diffusion and Afterburning (DAB) and Simultaneous Mixing and Combustion (SMC) test conditions. Results from both the 2D and the 3D models are presented.

Holt, James B.↗
DOE OSTI2020

Significance of the Direct Excitation Mechanism for High-Frequency Response of Premixed Flames to Flow Oscillations

Premixed flames are sensitive to flow disturbances, which can arise from acoustic or vortical fluctuations. For transverse instabilities, it is known that a dominant mechanism for flame response is “injector coupling,” whereby pressure oscillations associated with transverse waves excite axial flow disturbances. These axial flow disturbances then excite heat release oscillations. The objective of this paper is to consider another mechanism—the direct sensitivity of the unsteady heat release to transverse acoustic waves—and to compare its significance relative to the induced axial disturbances, in a linear framework. The rate at which the flame adds energy to the disturbance field is quantified using the Rayleigh criterion and evaluated over a range of control parameters, such as flame length and swirl number. The results show that radial modes induce heat release fluctuations that always add energy to the acoustic field, whereas heat release fluctuations induced by mixed radial-azimuthal modes can add or remove energy. These amplification rates are then compared to the flame response from induced axial fluctuations. For combustor-centered flames, these results show that the direct excitation mechanism has negligible amplification rates relative to the induced axial mechanism for radial modes. For transverse modes, the fact that the nozzle is located at a pressure node indicates that negligible induced axial velocity disturbances are excited; as such, the direct mechanism dominates. As a result, for flames that are not centered on pressure nodes, the direct mechanism for mixed modes dominates for certain nozzle locations and flame angles.

42 ENGINEERING↗
NASA NTRS1981

Component research for future propulsion systems

A review of factors related to the acquisition and life-cycle cost, and mission reliability of helicopters is given. The potential for advanced vehicle configurations with improvements in energy efficiency, operating economics, and characteristics to satisfy the demands of the future market are identified. Special attention is given to advanced propulsion systems and related component technologies, and system requirements, powerplants and component thrusts, compressor designs, combustion systems, turbine efficiency, blade tip treatment concepts and shaft dynamics are discussed in detail.

Walker, C. L.↗
NASA NTRS1986

Physics through the 1990s: Plasmas and fluids

The volume contains recommendations for programs in, and government support of, plasma and fluid physics. Four broad areas are covered: the physics of fluids, general plasma physics, fusion, and space and astrophysical plasmas. In the first section, the accomplishments of fluid physics and a detailed review of its sub-fields, such as combustion, non-Newtonian fluids, turbulence, aerodynamics, and geophysical fluid dynamics, are described. The general plasma physics section deals with the wide scope of the theoretical concepts involved in plasma research, and with the machines; intense beam systems, collective and laser-driven accelerators, and the associated diagnostics. The section on the fusion plasma research program examines confinement and heating systems, such as Tokamaks, magnetic mirrors, and inertial-confinement systems, and several others. Finally, theory and experiment in space and astrophysical plasma research is detailed, ranging from the laboratory to the solar system and beyond. A glossary is included.

Source record↗
NASA NTRS1990

Control-oriented modeling of combustion and flow processes in liquid propellant rocket engines

This paper presents a control-oriented model of the flow, reaction, and transport processes in liquid propellant rocket combustion chambers, based on the multicomponent conservation laws of gas dynamics. This model provides a framework for the inclusion of detailed chemical kinetic relations, viscous and other dissipative effects, a variety of actuators and sensors, as well as process and measurement disturbances. In addition to its potential usefulness to the designer in understanding the dynamical complexity of the system and the sources of model uncertainty, the model provides a rigorous basis for control system design. An appraisal of current and feasible actuators and sensors, and their mathematical representation are included.

Bentsman, Joseph↗
NASA NTRS2005

Low-Speed Active Flow Control Laboratory Developed

The future of aviation propulsion systems is increasingly focused on the application of control technologies to significantly enhance the performance of a new generation of air vehicles. Active flow control refers to a set of technologies that manipulate the flow of air and combustion gases deep within the confines of an engine to dynamically alter its performance during flight. By employing active flow control, designers can create engines that are significantly lighter, are more fuel efficient, and produce lower emissions. In addition, the operating range of an engine can be extended, yielding safer transportation systems. The realization of these future propulsion systems requires the collaborative development of many base technologies to achieve intelligent, embedded control at the engine locations where it will be most effective. NASA Glenn Research Center s Controls and Dynamics Technology Branch has developed a state-of-the-art low-speed Active Flow Control Laboratory in which emerging technologies can be integrated and explored in a flexible, low-cost environment. The facility allows the most promising developments to be prescreened and optimized before being tested on higher fidelity platforms, thereby reducing the cost of experimentation and improving research effectiveness.

Culley, Dennis E.↗
NASA NTRS2012

Design and Analysis of an Isokinetic Sampling Probe for Submicron Particle Measurements at High Altitude

An isokinetic dilution probe has been designed with the aid of computational fluid dynamics to sample sub-micron particles emitted from aviation combustion sources. The intended operational range includes standard day atmospheric conditions up to 40,000-ft. With dry nitrogen as the diluent, the probe is intended to minimize losses from particle microphysics and transport while rapidly quenching chemical kinetics. Initial results indicate that the Mach number ratio of the aerosol sample and dilution streams in the mixing region is an important factor for successful operation. Flow rate through the probe tip was found to be highly sensitive to the static pressure at the probe exit. Particle losses through the system were estimated to be on the order of 50% with minimal change in the overall particle size distribution apparent. Following design refinement, experimental testing and validation will be conducted in the Particle Aerosol Laboratory, a research facility located at the NASA Glenn Research Center to study the evolution of aviation emissions at lower stratospheric conditions. Particle size distributions and number densities from various combustion sources will be used to better understand particle-phase microphysics, plume chemistry, evolution to cirrus, and environmental impacts of aviation.

Heath, Christopher M.↗
NASA NTRSDate not supplied

CFD Evaluation of Aviation Fuels for Commercial Supersonics Technology

An overview is provided of a CFD analysis of a Lean-Direct Injection (LDI) combustion concept for NASA’s Commercial Supersonic Transport (CST) program. The National Combustion Code (OpenNCC) was used to perform two-phase reacting flow computations with three different aviation fuels for UTRC’s Axially Controlled Stoichiometry (ACS) combustor at CST cruise conditions. All computations were performed with a consistent approach of mesh-generation, spray and chemical-kinetics modeling for the combustor. Emissions (EICO, EINOx) and performance characteristics were predicted with OpenNCC, to evaluate the impact of replacing Jet-A with RP2 and Iso-Paraffinic Kerosene (IPK) for supersonic combustor applications. The CFD predictions of flame structure for IPK fuel were significantly different in comparison to those of Jet-A and RP2 fuel. However, the CFD analysis predicted very similar EINOx emissions for Jet-A, RP-2 and IPK fuels. The predicted EINOx emissions for all three fuels also compared very well with measured experimental data.

gas turbine combustion↗
NASA NTRS1998

Dynamics of Variable Mass Systems

This report presents the results of an investigation of the effects of mass loss on the attitude behavior of spinning bodies in flight. The principal goal is to determine whether there are circumstances under which the motion of variable mass systems can become unstable in the sense that their transverse angular velocities become unbounded. Obviously, results from a study of this kind would find immediate application in the aerospace field. The first part of this study features a complete and mathematically rigorous derivation of a set of equations that govern both the translational and rotational motions of general variable mass systems. The remainder of the study is then devoted to the application of the equations obtained to a systematic investigation of the effect of various mass loss scenarios on the dynamics of increasingly complex models of variable mass systems. It is found that mass loss can have a major impact on the dynamics of mechanical systems, including a possible change in the systems stability picture. Factors such as nozzle geometry, combustion chamber geometry, propellant's initial shape, size and relative mass, and propellant location can all have important influences on the system's dynamic behavior. The relative importance of these parameters on-system motion are quantified in a way that is useful for design purposes.

Eke, Fidelis O.↗
NASA NTRS1997

A Dual-Line Detection Rayleigh Scattering Diagnostic Technique for the Combustion of Hydrocarbon Fuels and Filtered UV Rayleigh Scattering for Gas Velocity Measurements

Non-intrusive techniques for the dynamic measurement of gas flow properties such as density, temperature and velocity, are needed in the research leading to the development of new generation high-speed aircraft. Accurate velocity, temperature and density data obtained in ground testing and in-flight measurements can help understand the flow physics leading to transition and turbulence in supersonic, high-altitude flight. Such non-intrusive measurement techniques can also be used to study combustion processes of hydrocarbon fuels in aircraft engines. Reliable, time and space resolved temperature measurements in various combustor configurations can lead to a better understanding of high temperature chemical reaction dynamics thus leading to improved modeling and better prediction of such flows. In view of this, a research program was initiated at Polytechnic University's Aerodynamics Laboratory with support from NASA Lewis Research Center through grants NAG3-1301 and NAG3-1690. The overall objective of this program has been to develop laser-based, non-contact, space- and time-resolved temperature and velocity measurement techniques. In the initial phase of the program a ND:YAG laser-based dual-line Rayleigh scattering technique was developed and tested for the accurate measurement of gas temperature in the presence of background laser glare. Effort was next directed towards the development of a filtered, spectrally-resolved Rayleigh/Mie scattering technique with the objective of developing an interferometric method for time-frozen velocity measurements in high-speed flows utilizing the uv line of an ND:YAG laser and an appropriate molecular absorption filter. This effort included both a search for an appropriate filter material for the 266 nm laser line and the development and testing of several image processing techniques for the fast processing of Fabry-Perot images for velocity and temperature information. Finally, work was also carried out for the development of a new laser-based strain-rate and vorticity technique for the time-resolved measurement of vorticity and strain-rates in turbulent flows.

Otugen, M. Volkan↗
DOE OSTI2026

Characterization of Flashback and Flame-Holding in a Jet-in-Crossflow Mixing Configuration With Methane–Hydrogen Fuel Blends

An approach that combines experimental and numerical analyses has been implemented to characterize the fundamentals of flashback events and flame-holding phenomena during high-hydrogen combustion in a jet-in-crossflow (JICF) configuration. Such flame dynamics are visualized experimentally using nanosecond (ns)-based hydroxyl planar laser-induced fluorescence (OH-PLIF) and chemiluminescence diagnostics techniques. The JICF burner has an optically accessible premixing tube allowing the optical diagnostics. The testing was conducted for varied premixer velocities (V) and equivalence ratio (ϕ) for 90–100% (H 2 , by mole) H 2 /CH 4 reactant mixtures at atmospheric temperature and pressure conditions. Two distinct flashback events were identified – conventional rich flashback and lean flashback, recorded while increasing ϕ and decreasing ϕ, respectively. The cause of lean flashback was attributed to lower momentum flux ratio which bends the jet sharply, closer toward the injection plane. The mean OH-PLIF images characterized the flame-holding behavior, where the flame was found to be stabilized on the leeward side only or on both windward and leeward sides as a lifted flame near the fuel port. A large eddy simulation (LES) with detailed chemistry combustion modeling approach was implemented along with an OH* submechanism, and it showed qualitative agreement with the integrated line-of-sight chemiluminescence results as well as the planar OH-PLIF measurements.

chemiluminescence diagnostic↗
DOE OSTI2025

Characterization of Flashback and Flame-Holding in a Jet-in-Crossflow Mixing Configuration with Methane-Hydrogen Fuel Blends

An approach that combines experimental and numerical analyses has been implemented to characterize the fundamentals of flashback events and flame-holding phenomena during high-hydrogen combustion in a jet-in-crossflow (JICF) configuration. Such flame dynamics are visualized experimentally using nanosecond (ns)-based hydroxyl planar laser-induced fluorescence (OH-PLIF) and chemiluminescence diagnostics techniques. The JICF burner has an optically accessible pre-mixing tube allowing the optical diagnostics. The testing was conducted for varied pre-mixer velocities (V) and equivalence ratio (ϕ) for 90%-100% (H2, by mole) H2/CH4 reactant mixtures at atmospheric temperature and pressure conditions. Two distinct flashback events were identified – conventional rich flashback and lean flashback, recorded while increasing ϕ and decreasing ϕ, respectively. The cause of lean flashback was attributed to lower momentum flux ratio which bends the jet sharply, closer toward the injection plane. The mean OH-PLIF images characterized the flame-holding behavior where the flame was found to be stabilized on the leeward side only or on both windward and leeward sides as a lifted flame near the fuel port. A large eddy simulation (LES) with detailed chemistry combustion modeling approach was implemented along with an OH* sub-mechanism, and it showed qualitative agreement with the integrated line-of-sight chemiluminescence results as well as the planar OH-PLIF measurements.

chemiluminescence diagnostic↗