COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS

1-D simulations based on the quasi-one-dimensional equations of fluid motion plus an ignition delay model and 2-D numerical simulations based on Reynolds-Averaged Navier-Stokes (RANS) equations have been performed for two different scramjet combustors. The combustor configurations at DLR and NASA's SCHOLAR Supersonic Combustor have been used as test cases for the 1-D and 2-D simulations. Comparisons between the published 3-D computational and experimental results and quasi-one-dimensional and 2-D simulations have been performed. The quasi-one dimensional modeling of NASA's SCHOLAR supersonic combustor captures the trends in Mach number, static pressure and static temperature for both cold flow and combustion case. The comparison with experimental result for combustion case reveals a close agreement with the pressure peak and the presence of an ignition delay. Thus, 1-D simulation very closely predicts the flow evolution within the combustor. On the other hand, for DLR supersonic combustor, due to the lack of oblique wave (i.e. shock waves and expansion waves) and shear dominated viscous flow simulation, 1-D model severely fails to predict the trend followed by the experimental result along the centerline of the combustor. However, the 1-D model is able to match the overall flow velocity achieved within the combustor downstream of the wedge at approximately six wedge chord lengths

Langley aerospace test highlights, 1985

The role of the Langley Research Center is to perform basic and applied research necessary for the advancement of aeronautics and space flight, to generate new and advanced concepts for the accomplishment of related national goals, and to provide research advice, technological support, and assistance to other NASA installations, other government agencies, and industry. Significant tests which were performed during calendar year 1985 in Langley test facilities, are highlighted. Both the broad range of the research and technology activities at the Langley Research Center and the contributions of this work toward maintaining United States leadership in aeronautics and space research, are illustrated. Other highlights of Langley research and technology for 1985 are described in Research and Technology-1985 Annual Report of the Langley Research Center

Pseudo-shock waves and their interactions in high-speed intakes

In an air-breathing engine the flow deceleration from supersonic to subsonic conditions takes places inside the isolator through a gradual compression consisting of a series of shock waves. The wave system, referred to as a pseudo-shock wave or shock train, establishes the combustion chamber entrance conditions, and therefore influences the performance of the entire propulsion system. The characteristics of the pseudo-shock depend on a number of variables which make this flow phenomenon particularly challenging to be analysed. Difficulties in experimentally obtaining accurate flow quantities at high speeds and discrepancies of numerical approaches with measured data have been readily reported. Understanding the flow physics in the presence of the interaction of numerous shock waves with the boundary layer in internal flows is essential to developing methods and control strategies. To counteract the negative effects of shock wave/boundary layer interactions, which are responsible for the engine unstart process, multiple flow control methodologies have been proposed. Improved analytical models, advanced experimental methodologies and numerical simulations have allowed a more in-depth analysis of the flow physics. The present paper aims to bring together the main results, on the shock train structure and its associated phenomena inside isolators, studied using the aforementioned tools. Several promising flow control techniques that have more recently been applied to manipulate the shock wave/boundary layer interaction are also examined in this review

Airframe-Propulsion Integration Design and Optimization

Airframe-propulsion integration design is one of the key technologies of the hypersonic vehicle. With the development of hypersonic vehicle design method, CFD technology, and optimization method, it is possible to improve the conceptual design of airframe-propulsion integration both in accuracy and efficiency. In this chapter, design methods of waverider airframes and propulsion systems, including inlets, nozzles, isolators, and combustors, are reviewed and discussed in the light of CFD analyses. Thereafter, the Busemann inlet, a three-dimensional flow-stream traced nozzle, and a circular combustor together with a cone-derived waverider are chosen to demonstrate the airframe-propulsion integration design. The propulsion system is optimized according to the overall performance, and then the component such as the nozzle is optimized to obtain a better conceptual configuration

Numerical Investigation of Hydrogen-fuelled Scramjet Combustor with Cavity Flame Holder

Reacting flow field of a cavity-based hydrogen-fuelled supersonic combustion ramjet (scramjet) combustor has been explored numerically. 3-D RANS equations are solved alongwith k-ε turbulence model and infinitely fast rate kinetics for combustion. The flow field inside the flame holding cavity and its effect on the mixing and reaction are explored in detail by performing simulations of two different combustors (with and without cavities) for which elaborate surface pressure measurements are available for reacting and non-reacting flows. Simulations capture all the finer details of the flow field and good match between computed surface pressures experimental values for different equivalence ratios forms the basis of further analysis. The cavity of the combustor behaves as an open cavity for both non-reacting and reacting flow-even though the flow patterns inside the cavity are quite different for both the cases. Flame-holding cavities are seen to augment mixing and reaction in small sized combustors.Defence Science Journal, Vol. 64, No. 5, September 2014, pp.417-425, DOI:http://dx.doi.org/10.14429/dsj.64.519

Performance Measurements of Direct Air Injection in a Cavity-Based Flameholder for a Supersonic Combustor

For several years the Air Force Research Lab Propulsion Directorate has been studying the difficulties in fueling supersonic combustion ramjet engines with hydrocarbon based fuels. Recent investigations have focused on the use of direct air injection into a directly-fueled cavity-based flameholder. Direct air injection has been shown qualitatively to be a valuable tool for improving cavity combustion. Little quantitative data is available that characterizes the performance of cavity-based flameholders. The objective of this research was to quantitatively determine the specific advantages and disadvantages of the direct air injection scheme. This was accomplished via intrusive probing into a supersonic free stream flow at an axial location behind the cavity flameholder. Pitot and static pressure, total temperature, and gas sampling measurements were taken and the corresponding values were processed to yield relevant engineering quantities. Data were taken over a range of fuel and air injection rates. Direct air injection resulted in increased combustion throughout the area of interest behind the cavity. Air injection increased the static temperature and pressure throughout the area of interest. Enthalpy spread into the free stream was also increased through the use of air injection. Total pressure losses increased as a result of the direct air injection scheme. The ratio of enthalpy increase to increase in total pressure losses increased with higher fuel flow rates, indicating that the direct air injection technique shows more promise for higher fuel loadings

Three-Dimensional Wall Effects in a Scramjet Cavity Flameholder

Scramjets, or supersonic combustion ramjets, are a key component for the development of powered hypersonic vehicles. The flameholding strategy used in this engine creates a recirculation region by creating a cavity in one side of the combustor. This highly unsteady region is difficult to interrogate through conventional means. A novel optical observation technique called hyperspectral imaging has been developed to examine the scramjet combustion chamber. The hyperspectral camera is capable of generating an interferogram at hundreds to thousands of wavelengths. These data are integrated across the line-of-sight with no information on three-dimensional (3D) location of origin. A model must be used to extrapolate spatially-resolved two dimensional (2D) data to a three dimensional domain. With no a priori data to inform otherwise, the current model assumes that the scramjet flowfield is uniform in the spanwise direction. This does not agree with understanding of compressible ow theory of shock dominated and turbulent flows. This work simulates a non-reacting scramjet combustor using hybrid Reynolds-Averaged Navier Stokes (RANS)/Large Eddy Simulation (LES). The spanwise character is analyzed through instantaneous and time-averaged statistics. It is shown that the flow is not well approximated as two-dimensional, especially in the cavity where fuel is transported and mixed with air. Complete time histories of spanwise lines were collected and total time-averaged means were compared against 200 µs windowed means. These time windows correspond with the time it takes the hyperspectral camera to create a single scan. Turbulent time scales are calculated and their ramifications on the collection process of the hyperspectral camera are considered. The viscous-dampened region in the cavity has integral timescales on the order of the collection time of each interferogram

Study of hypersonic propulsion/airframe integration technology

An assessment is done of current and potential ground facilities, and analysis and flight test techniques for establishing a hypersonic propulsion/airframe integration technology base. A mach 6 cruise prototype aircraft incorporating integrated Scramjet engines was considered the baseline configuration, and the assessment focused on the aerodynamic and configuration aspects of the integration technology. The study describes the key technology milestones that must be met to permit a decision on development of a prototype vehicle, and defines risk levels for these milestones. Capabilities and limitations of analysis techniques, current and potential ground test facilities, and flight test techniques are described in terms of the milestones and risk levels

Aeronautical engineering: A continuing bibliography with indexes (supplement 321)

This bibliography lists 496 reports, articles, and other documents introduced into the NASA scientific and technical information system in Sep. 1995. Subject coverage includes: design, construction and testing of aircraft and aircraft engines; aircraft components, equipment, and systems; ground support systems; and theoretical and applied aspects of aerodynamics and general fluid dynamics