Skin Friction Reduction in Hypersonic Turbulent Flow by Boundary Layer Combustion

Suraweera, Milinda V.; Mee, D. J.; Stalker, R. J.

doi:10.2514/6.2005-613

Cited by 37 publications

(24 citation statements)

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“…In the combustion region downstream the jet, the eddy viscosity is about 100 times to 150 times the molecular viscosity, and the diffusion of the species and temperature is enhanced greatly. It is also seen from Figure 8 that downstream the jet, the eddy viscosity near the upper wall is greater than that near the lower wall, which illuminates an important effect of combustion on the development of turbulence [16,17]: the heat release of combustion induces the expansion of the fluid volume and the decrease of average density, which can reduce the production of Reynolds stress and then the eddy viscosity in turbulence model. Figure 10 shows the vortex structure near the jet hole.…”

Section: Analysis Of Fluid Characteristicsmentioning

confidence: 88%

A numerical study of turbulent combustion characteristics in a combustion chamber of a scramjet engine

Gao

Lee

2010

Sci. China Technol. Sci.

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3D numerical simulation of flow fields in a combustion chamber of a scramjet engine using an SST turbulence model with an explicit compressibility correction was performed and the results were compared to the experimental results. The characteristics of the turbulent combustion flow fields were analyzed via the numerical results and presented. In order to identify the mechanisms of turbulent combustion in supersonic flows, the evolutions of governing dimensionless parameters in the flow fields were investigated based on the theory of combustion and the available numerical results. It was found that the supersonic combustion takes place in the region of fully developed turbulence and that the strongest effects of turbulence and combustion processes appear in the vicinity of the injector. The unsteady effects and the local flame extinction phenomenon induced by turbulent flows were found to be negligibly small, and the steady flamelet approximation will hold for practical applications. scramjet, turbulent combustion, flamelet approximation, numerical simulation Citation:Gao Z X, Lee C H. A numerical study of turbulent combustion characteristics in a combustion chamber of a scramjet engine.

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Section: Analysis Of Fluid Characteristicsmentioning

confidence: 88%

A numerical study of turbulent combustion characteristics in a combustion chamber of a scramjet engine

Gao

Lee

2010

Sci. China Technol. Sci.

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“…The main objective of this particular injection configuration is to reduce skin-friction. Suraweera (2006) showed that skin-friction drag in a scramjet combustor could be reduced by up to 80% by burning hydrogen within the boundary layer. This injection scheme is important to the operation of the RESTM12 scramjet engine since at the Mach 12 design point drag will be extremely high any potential method of reducing this must be investigated.…”

Section: Step Injectionmentioning

confidence: 99%

Experimental investigation of a three dimensional scramjet engine at hypervelocity conditions

Wise¹

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Air-breathing propulsion has the ability to provide more economical access-to-space than current rocket based systems. Access-to-space requires hypersonic flight within the atmosphere, the air-breathing cycle best suited to this is the scramjet. Next generation launch systems require a scramjet stage to operate into the hypervelocity regime (> 3 km/s). Performance data at these conditions is scarce due to the inability of ground test facilities to replicate the high total pressures and temperatures associated with flight. Also, flight experiments are subject to material limitations and are generally prohibitively expensive. This thesis details experiments which add to the currently limited data-set of scramjet performance at hypervelocity conditions.The Mach 12 rectangular-to-elliptical shape-transitioning (RESTM12) engine is currently under investigation at The University of Queensland as a potential candidate for access-to-space. The RESTM12 engine is designed to accelerate from Mach 6-12 as a part of a rocket-scramjet-rocket access-to-space system. Previous testing of the RESTM12 engine has shown good performance at off-design conditions and now as a result of this investigation, robust performance at its design condition.The primary aim of this thesis was to:"investigate experimentally whether robust supersonic combustion can be generated in a three-dimensional scramjet flowpath at a Mach 12 flight condition in an impulse facility."In order to achieve this two experimental campaigns where undertaken in the T4 Stalker Tube. Both experiments used the facilities Mach 10 nozzle to generate a test flow equivalent to Mach 11.8 flight at 38.3 km altitude. Initial experiments investigating hypervelocity boundary layer transition showed that natural transition would not occur in a length equivalent to the forebody and inlet of the engine. The ingestion of a laminar boundary layer into the inlet of the RESTM12 engine would increase the likelihood of developing a separation in the inlet that would not allow it to operate as intended. To reduce the susceptibility of the inlet to separation a boundary layer trip was required.i Based on the Hyper-X program's boundary layer trip development experiments (Berry et al., 2001a), several combinations of trip geometry, height and location were tested on a flat plate model resulting in a successful trip configuration. Now confident that a trip configuration capable of providing the RESTM12 engine with a turbulent boundary layer was designed, the engine was tested at its design condition.A half-scale model of the RESTM12 scramjet engine was tested in a semi-free-jet configuration utilising both inlet and step injection of gaseous hydrogen fuel. Once it was established that the inlet was operating correctly and sufficient test time was available the experiments began with fuel injection. Initial tests showed that robust combustion occurred when inlet injection was employed. The most successful fuelling configuration was a combined injection scheme where 31% of the fuel was ...

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“…Actually, Goyne et al [9] and Suraweera et al [10] measured the Stanton number with and without boundary-layer combustion in their experiments, and they found that the Stanton number, rather than being increased under the heat-release effects, was slightly reduced. However, for this somewhat surprising result, no physical explanation was given.…”

mentioning

confidence: 97%

“…Boundary-layer combustion is then considered as an effective skin-friction reduction technique for scramjet engines and draws continuous attention. In 2005, Suraweera et al [10] conducted further experimental research on boundary-layer combustion to examine whether it is still effective for viscous drag reduction at different flight speeds. Their experimental data proved that skinfriction reductions by boundary-layer combustion were achieved at all values of the freestream total enthalpy in the experiments.…”

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confidence: 98%

Combustion Heat-Release Effects on Supersonic Compressible Turbulent Boundary Layers

et al. 2015

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Influences of heat release by the hydrogen combustion in supersonic turbulent boundary layers are numerically studied using Reynolds-averaged Navier-Stokes equations. The adopted Reynolds-averaged Navier-Stokes methodology is first validated by comparing the numerical results with the existing experimental data. Studies on the effects of the flame perpendicular position inside the boundary layer reveal that, while the flame is restricted around the edge of the boundary layer, the heat release may slightly reduce rather than increase the wall heat flux because of the suppression effect on the turbulent energy transport due to heat release. However, as the flame moves toward the wall, the skin-friction reduction effect would not be obviously strengthened, but the wall heat flux could be dramatically enhanced by the increase of near-wall chemical reactions. At a given hydrogen mass flow rate, the injection scheme with a higher injection height and a lower injection velocity could be helpful to achieve a larger skinfriction reduction while maintaining a lower wall heat flux. Finally, analysis of the heat-release effects on the velocity law of the wall shows that van Driest's velocity law largely deviates from the computational results, whereas White's velocity law remains close to the numerical results within a region of approximately y < 300. Nomenclature B = law-of-the-wall constant (5.0) c f = local skin friction coefficient c p = specific heat at constant pressure, J∕kg · K D = mass diffusivity, m 2 ∕s E = total energy, J∕kg H = total enthalpy, J∕kg h s = static enthalpy for species s, J∕kg h = hydrogen injection height at the entrance, m k = turbulent kinetic energy, m 2 ∕s 2 M = Mach number Pr = Prandtl number Pr t = turbulent Prandtl number p = pressure, MPa q w = local heat flux, MW∕m 2 r= Pr 1∕3 , recovery factor Sc t = turbulent Schmidt number T = temperature, K u = velocity, m∕s u τ = friction velocity, m∕s Y hw = hydrogen mass fraction at the wall Y s = mass fraction of species s δ = boundary-layer thickness, m κ = von Kármán constant (0.41) λ = molecular thermal conductivity, W∕m · K μ = molecular viscosity, N · s∕m 2 μ t = turbulent eddy viscosity, N · s∕m 2 ρ = density, kg∕m 3 τ = shear stress, N∕m 2 ω = specific turbulent dissipation, 1∕s _ ω = mass production per unit volume per unit time, kg∕m 3 · s Subscripts e = edge of boundary layer j = jet s = species s w = wall ∞ = freestream Superscripts t = turbulent = Reynolds averagẽ = Favrè average

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Skin Friction Reduction in Hypersonic Turbulent Flow by Boundary Layer Combustion

Cited by 37 publications

References 10 publications

A numerical study of turbulent combustion characteristics in a combustion chamber of a scramjet engine

A numerical study of turbulent combustion characteristics in a combustion chamber of a scramjet engine

Experimental investigation of a three dimensional scramjet engine at hypervelocity conditions

Combustion Heat-Release Effects on Supersonic Compressible Turbulent Boundary Layers

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