An experimental investigation of the cold flowfield in a model scramjet combustor

Zhao, Zihao; Song, Wenping; Xiao, Yinli; Le, Jialing

doi:10.1243/09544100jaero479

Cited by 7 publications

(4 citation statements)

References 15 publications

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“…To solve this problem, some fuel injection techniques have been proposed. The strut [5][6][7][8][9][10][11][12][13][14], the cavity [15][16][17][18], the cantilevered ramp [4], the backward facing step [19][20][21] and the combination [22][23][24][25], whose principle is to generate vorticity in the vicinity of the walls of the combustor. The region forming the vorticity, namely the recirculation zone, has low flow velocity.…”

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confidence: 99%

Parametric effects on the combustion flow field of a typical strut-based scramjet combustor

et al. 2011

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The flame-holding mechanism in hypersonic propulsion technology is the most important factor in prolonging the duration time of hypersonic vehicles. The two-dimensional coupled implicit Reynolds-averaged Navier-Stokes equations, the shear-stress transport k- turbulence model and the finite-rate/eddy-dissipation reaction models were used to simulate the combustion flow field of a typical strut-based scramjet combustor. We investigated the effects of the hydrogen-air reaction mechanism and fuel injection temperature and pressure on the parametric distributions in the combustor. The numerical results show qualitative agreement with the experimental data. The hydrogen-air reaction mechanism makes only a slight difference in parametric distributions along the walls of the combustor, and the expansion waves and shock waves exist in the combustor simultaneously. Furthermore, the expansion wave is formed ahead of the shock wave. A transition occurs from the shock wave to the normal shock wave when the injection pressure or temperature increases, and the reaction zone becomes broader. When the injection pressure and temperature both increase, the waves are pushed out of the combustor with subsonic flows. When the waves are generated ahead of the strut, the separation zone is formed in double near the walls of the combustor because of the interaction of the shock wave and the boundary layer. The separation zone becomes smaller and disappears with the disappearance of the shock wave. Because of the horizontal fuel injection, the vorticity is generated near the base face of the strut, and this region is the main origin for turbulent combustion. , hypersonic propulsion technology has drawn increasing attention from researchers. However, the flame-holding mechanism in the scramjet combustor cannot satisfy the requirement for cruising a long time in near-space [3] because of the short residence time of the mixture remaining in the supersonic flow, which is in the order of milliseconds [4]. This restricts improvement of the aerodynamic performance of hypersonic vehicles. To solve this problem, some fuel injection techniques have been proposed. The strut [5][6][7][8][9][10][11][12][13][14], the cavity [15][16][17][18], the cantilevered ramp [4], the backward facing step [19][20][21] and the combination [22][23][24][25], whose principle is to generate vorticity in the vicinity of the walls of the combustor. The region forming the vorticity, namely the recirculation zone, has low flow velocity. In this region the fuel mixes with the supersonic flow and the mixture can stay in the scramjet combustor for a long time. Zou et al. [6] have used a newly-proposed partiallyresolved numerical simulation procedure to investigate turbulent combustion in a two-dimensional DLR scramjet engine, and the large-scale turbulence was defined by the temporal filtering. Oevermann [7] has used a two-equation k- turbulence model combined with a stretched laminar

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confidence: 99%

Parametric effects on the combustion flow field of a typical strut-based scramjet combustor

et al. 2011

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“…8,9 Several investigations have also confirmed the correlation between natural shock, 10 the effect of roughness on the wall, 11 and increased surface heating. 12 In addition, some research has focused on improvements in scramjet intake, 13 turbine, 14 and compressor 15 based on SWBLI.…”

Section: Introductionmentioning

confidence: 99%

A computational study of leading-edge wall cooling effect on shock wave–boundary layer interaction in forward-facing step

Murugan

2023

Proceedings of the Institution of Mechanical Engineers, Part G:

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A computational study has been done to understand the effect of leading-edge wall cooling on shock wave–boundary layer interaction. Shock wave–boundary layer interaction is studied over a forward-facing step at supersonic Mach 2.5. The study was carried out using Ansys. The work aims to explore the implementation of wall cooling at the leading edge as a separation control strategy for supersonic forward-facing step-induced flow separation. We use a finite-volume method based on upwind flux difference splitting and second-order upwind flow discretization. The simulation results are validated with the available experimental data. Furthermore, using numerical simulations, we found that the separation bubble size was reduced by 18.36% when the walls were marginally cooled to 150 K, while the separation was reduced by 32.65% when the walls were strongly cooled to 100 K.

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“…This interaction may lead to undesirable effects in the vicinity of interaction such as boundary-layer separations 1,2 and higher thermal loads. 3,4 To mitigate these issues, many active 5,6 as well as passive 7,8 techniques were proposed by various researchers.…”

Section: Introductionmentioning

confidence: 99%

Comparative studies of shock-wave boundary-layer interactions in Earth and Mars atmospheres

Das

Desai

Kulkarni

2021

Proceedings of the Institution of Mechanical Engineers, Part G:

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Investigations of ramp-induced shock-wave boundary-layer interaction have been carried out for real gas flows of air and carbon dioxide through hypersonic laminar flow simulations corresponding to Earth and Mars atmospheres. An in-house-developed solver, which accounts for the real gas effects, has been employed for these studies. Effects of various parameters like wall temperature, freestream stagnation enthalpy, freestream Mach number, and blunt leading edge are explored on the intensity of shock-wave boundary-layer interaction (SWBLI). In either case, an increase in separation length is observed with an increase in wall temperature and a decrease in Mach number as well as freestream stagnation enthalpy. Here, the intensity of alteration is always noted to have a higher percentage for the Mars gas model. Further, separation length is found to be almost equal for the same wall to total temperature ratio in both of the flow mediums. The present study also affirms the fact that the leading edge bluntness can be used as a tool to reduce the size of the separation region in these planetary atmospheres. Revised correlations have been proposed for hypersonic Earth atmospheric flow with real gas effects to predict the extent of upstream influence and separation bubble size. The outcomes of simulations have also helped to device new correlations for these flow features of SWBLI for Mars atmospheric conditions. In all, the need for consideration of real gas effects and an exclusive real gas flow solver for the Mars atmosphere are the prominent recommendations of current studies.

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An experimental investigation of the cold flowfield in a model scramjet combustor

Cited by 7 publications

References 15 publications

Parametric effects on the combustion flow field of a typical strut-based scramjet combustor

Parametric effects on the combustion flow field of a typical strut-based scramjet combustor

A computational study of leading-edge wall cooling effect on shock wave–boundary layer interaction in forward-facing step

Comparative studies of shock-wave boundary-layer interactions in Earth and Mars atmospheres

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