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Smart M.K.,University of Queensland | Smart M.K.,Center for Hypersonics
AIAA Journal | Year: 2015

Scramjet operation in the lower hypersonic regime between Mach 4 and 7 is characterized by what is called dualmode combustion. In this situation, the backpressure generated by heat release in the combustor separates the wall boundary layer upstream of fuel injection, forming a highly complex flow known as a pseudoshock. The pseudoshock phenomenon has been studied extensively in the experimental literature, and its structure has been found to be affected by the duct Mach number, the extent of the wall boundary layer, and the streamwise area distribution of the duct. This paper presents a quasi-one-dimensional analysis technique for calculating the flow through a pseudoshock. It is quite general, and it can be used for ducts with any supersonic inflow, cross-sectional shape, and streamwise area distribution. Comparison of the results of this analysis with experimental data indicates that it provides a good estimate of both the interaction length and the pressure distribution forX-type pseudoshocks that occur in ducts with Mach numbers above two. © 2015 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc. Source


Smart M.K.,University of Queensland | Smart M.K.,Center for Hypersonics
AIAA Journal | Year: 2012

The supersonic combustion ramjet, or scramjet, is the engine cycle most suitable for sustained hypersonic flight in the atmosphere. This paper examines a key decision in the design of the inlet or intake of these hypersonic airbreathing engines, namely, the level of compression to be performed. Too much compression can lead to onerous system level issues including the need for bleed or variable geometry, while too little compression can result in low cycle efficiency and poor combustion of fuel. An analysis of the important factors that affect the choice of scramjet inlet compression ratio has been performed for hydrogen-fueled scramjets at Mach 6, 8, 10, and 12. It was found that contrary to classical thermodynamic analyses, scramjet cycle efficiency reaches an optimum at a relatively low compression ratio between 50 and 100 for all Mach numbers. Practical constraints related to nonequilibrium flow effects, inlet starting, and boundary-layer separation were also shown to prompt a desire for low compression ratio. The lower limit on compression was found to be set by the need to complete the combustion reaction in the available engine length and is therefore dependent on engine scale. On the basis of these factors it is recommended that scramjet inlet compression ratio be set to the minimum that satisfies the robust combustion requirement, with the caveat that it not be below 50 in order to maintain high cycle efficiency. For typical wind-tunnel-scale engines, this results in a requirement for the inlet to compress airflow entering the combustor to a pressure of approximately 1 2 atm, regardless of the flight Mach number. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Source


Moule Y.,University of Queensland | Smart M.K.,Center for Hypersonics
Journal of Propulsion and Power | Year: 2013

A study was conducted for performance analysis of a mach 12 scramjet being developed at the Center for Hypersonics, University of Queensland, Australia. A three-dimensional (3-D) scramjet, with a rectangular-to-elliptical shape transition (REST) inlet and an elliptical combustor was developed with a design point of Mach 12 for the investigations. The inlet injection and an injector scheme, which were designed to generate combustion in the boundary layer, were incorporated into this flowpath. Ground testing of this engine, with hydrogen fuel at off-design conditions simulating flight at Mach 8.7 showed robust combustion and good pressure rise from the two injection schemes. The inlet compression was determined through numerical simulation using the NASA Langley code, VULCAN for the current performance analysis. The goal of the performance analysis was to estimate the combustion efficiency curve for the different fueling configurations tested in the experiments. Source


Pudsey A.S.,University of Queensland | Pudsey A.S.,Center for Hypersonics | Boyce R.R.,Center for Hypersonics
Journal of Propulsion and Power | Year: 2010

A three-dimensional numerical study has been performed of the effects of sonic gaseous hydrogen injection through multiple transverse injectors subjected to a supersonic crossflow. Solutions were obtained for a series of injection configurations in a Mach 4.0 crossflow, with a global equivalence ratio of ø = 0:5. Results indicate a different flow structure than for a typical single jet, with the development of two clearly defined wake vortices, including a stagnation point and reversed flow region immediately behind each downstream jet. While the overall penetration was reduced under the investigated conditions, significant improvements were observed when nondimensionalizing against the equivalent jet diameter for each modeled injector row. This was found to be the result of increased jet-to-freestream momentum ratio due to the subsonic flow regions between each injector. Further enhancements were also observed in terms of mixing performance for the multijet cases. Improvements of up to 5% in the overall mixing efficiency were experienced by using multiple jets due to increased mixant interface area and intermediate stirring through wake vortices between each injector. No improvement in far-field mixing was observed. Overall, it has been demonstrated that there are benefits to be gained through the injection of gaseous hydrogen from many small injectors rather than fewer large injectors. Copyright © 2010 by A. S. Pudsey and R. R. Boyce. Published by the American Institute of Aeronautics and Astronautics, Inc. Source


Kang S.H.,Korea Aerospace Research Institute | Lee Y.J.,Korea Aerospace Research Institute | Yang S.S.,Korea Aerospace Research Institute | Michael K. Smart,University of Queensland | And 4 more authors.
Journal of Propulsion and Power | Year: 2011

To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the T4 free-piston shock tunnel. The test model had a rectangular intake, which compressed the freestream flow through a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of 31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40. Furthermore, the cavity flame holder and the W-shaped cowl showed improved pressure distribution due to greater reaction intensity. With the aid of numerical analysis, the cavity and the W-shaped cowl are shown to be effective in fuel-air mixing. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Source

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