Investigation On The Growth Characteristics And Heat Release Features Of The Supersonic Reacting Mixing Layer In Confined Space

With the development of hypersonic flight technology,turbulent combustion in high-speed flows,which is a basic subject of combustion science,has become a hot topic in the field of aerospace.And the key issue is the interaction of turbulence-combustion-shocks under the geometric constraints.In the present study,the hypersonic propulsion system was taken as the engineering background,and the main characteristics of the Rocket-Based Combined Cycle(RBCC)engine were extracted.Then the supersonic reacting mixing layer in the confined space was used to model the turbulent combustion between the supersonic rocket jet and the supersonic airstream.The focus was placed on the following two issues: the growth characteristics of the supersonic reacting mixing layer in the confined space affected by complex wave structures,and the heat release features in the compound flow field affected by multiple combustion modes.Large eddy simulation,shadow photography,tunable diode laser absorption spectroscopy,spontaneous radiation,and pressure sensors were used to study the supersonic reacting mixing layer in the confined space.According to the plots of the layer's thickness and the convective Mach number,the shocks' impacts on the compressibility were discussed.Both positive and negative effects of the shocks on the development of the layer were analyzed.Then the growth characteristics were obtained for the supersonic reacting mixing layer affected by shocks in the confined space.The diffusion and the premixed combustion modes along with the layer's growth were studied from the viewpoint of mixing patterns.And the supersonic and the subsonic combustion modes along with the layer's growth were investigated according to the flow types.The mapping relationship between the different combustion modes and the heat release rate was established through filtering functions.Then the released heat and the reaction zones for different combustion modes were calculated.Finally,several parameters such as the contribution coefficient,the ratio of the combustion area,and the modified-averaged Mach number in the flow channel were proposed to analyze the combustion modes and heat release features in the compound flow field.The results of this research indicate that:(1)Affected by shocks in the confined space,three typical regions in terms of the layer's thickness could be observed during the supersonic reacting mixing layer's growth: the initial region,the quickly-developing region,and the fully-developed region.The waves near the entrance of the combustor change the convective Mach number of the mixing layer rapidly,which makes the compressibility and the growth characteristics changed consequently.The shock waves in the channel have both negative and positive effects on the development of the reacting mixing layer.On the one hand,the compressive effect of the shock waves reduces the layer's thickness at the incident point,resulting in suppression for the layer's growth.On the other hand,the baroclinic effect of the shock waves increases the layer's thickness by inducing vorticity gain after the incident point,leading to the promotion for the layer's growth.Therefore,in the region where the shock waves interact with the reacting mixing layer,the layer's thickness decreases first and then increases in an oscillatory mode.The interaction between the shock and the layer is characterized as refraction or reflection according to the shock's intensity and the layer's gradients.(2)Local compression zones as well as expansion regions are formed in the flow field.The baroclinic effect associated with shocks could promote the generation of vortex,which is beneficial to the mixing and combustion between fuel and oxidizer.As a consequence,the heat release rate increases rapidly behind the interaction region.Then the gas expands in the local field and compresses the surrounding fluid to form the pneumatic compression surface.Besides,the intensity and the angle of the oblique shock in the interaction region increase with a tendency to form the normal shock.The significant heat release could enhance the transport process,reduce the flow gradients,and weaken the discontinuous features caused by shocks.Therefore,the attenuation of the shocks in the channel is accelerated and the wave structures in the downstream are damaged.(3)In a hypersonic propulsion system,the fuel jet and airstream are initially injected into the combustor in the non-premixed manner to form the reacting mixing layer.In the initial stage,the reaction between fuel and oxidizer is mainly featured as diffusion combustion due to the opposite gradients of the reactants.The mixing efficiency gradually increases as the layer grows in the channel.When the angle between fuel gradient and oxidizer gradient is acute,local premixed combustion zones would form in the flow field.Therefore diffusion combustion and premixed combustion take place simultaneously and contribute to heat release collectively.Shocks could promote the mixing between fuel and oxidizer due to the baroclinic effect,which could enhance the contributions of the premixed combustion mode to heat release.(4)Although both the fuel jet and the airstream flow into the chamber at supersonic speeds,local subsonic regions are still formed in the flow field when the flow speed is less than the local sonic speed due to geometry configuration,wave structure,vortex motion,and chemical reactions,etc.Therefore,supersonic combustion and subsonic combustion occur in the flow field at the same time and lead to heat release correspondingly.Although a little amount of heat is released in the initial region,it mainly results from the subsonic combustion mode because of the low-speed recirculation zone at the end of the central strut.The supersonic combustion mode plays a more and more important role on heat release when the recirculation zone disappears.Shocks are beneficial to subsonic combustion since they could reduce the Mach number in the local flow field.(5)Under the condition that the rocket's state is steady and the air flow rate is constant,the back pressure of the rocket nozzle rises with the increase of the incoming air temperature.Consequently,the expansion of the rocket jet is suppressed.Moreover,the length of the potential core decreases and the low-speed recirculation zone at the end of the central strut shrinks correspondingly.As a result,the reacting mixing layer enters the quickly-developing region in earlier phase and the inflection point in the layer's thickness curve moves toward the central strut.The increase of temperature for the incoming air enhances the reactivity of the reactants,making it easier to combust between the rocket jet and the airstream.And the amount of heat released in the combustor increases consequently.At the same time,it decreases the residence time for the high-speed flow in the chamber,which limits the rapid growth for the layer.Therefore,the combustion zone shrinks in the flow field.According to the mixing patterns,the diffusion combustion become more dominant since it releases more heat and corresponds to a larger reaction zone than the premixed combustion.Based on the flow types,the released heat and combustion zone associated with supersonic combustion increase while the ones regarded to subsonic combustion decreases,which illustrates that supersonic combustion plays a more and more important role in the chamber.