Martin Konopka

Large-eddy simulation of shock-cooling-film interaction at helium and hydrogen injection

Laminar helium and hydrogen films at a Mach number 1.3 are injected through a slot into a fully turbulent freestream air flow at a Mach number 2.44. To numerically study by large-eddy simulations the impact of an impinging shock on various cooling films, first, reference solutions without shock impingement are computed and then, the helium and hydrogen cooling films interacting with an oblique shock at a pressure ratio of p3/p1 = 2.5 are analyzed. The comparison of the helium and hydrogen injections without shock shows the hydrogen injection to have a 1.14-fold better cooling effectiveness at 60% of the blowing rate of the helium injection. The shock-cooling-film interaction causes a massive separation bubble that is 23% larger at the hydrogen than at the helium injection. Nevertheless, the shock influenced cooling effectiveness at the hydrogen injection is only 30% reduced compared to a 40% decrease at the helium injection 100 slot heights downstream of the injection. The intense mixing in the shock-cool...

Large-Eddy Simulation of Film Cooling in an Adverse Pressure Gradient Flow

In order to analyze the interaction of multiple rows of film cooling holes in flows at adverse pressure gradients, large-eddy simulations (LESs) are performed. The considered three-row cooling configuration consists of inclined cooling holes at an angle of 30 deg with a lateral pitch of p/D=3 and a streamwise spacing of l/D=6. The cooling holes possess a fan-shaped exit geometry with lateral and streamwise expansions. For each cooling row the complete internal flow is computed. Air and CO2 are injected in order to investigate the influence of an increased density ratio on the film cooling physics at adverse pressure gradients. The CO2 injected at the same blowing rate as air shows a higher magnitude of the Reynolds shear stress component and, thus, an enhanced mixing downstream of the cooling holes. The LES results of the air and CO2 configurations are compared to the corresponding particle-image velocimetry (PIV) measurements and show a convincing agreement in terms of the averaged streamwise velocity and streamwise velocity fluctuations. Furthermore, the cooling effectiveness is investigated for a zero and an adverse pressure gradient configuration with a temperature ratio at gas turbine conditions. For the adverse pressure gradient case, reduced temperature levels off the wall are observed. However, the cooling effectiveness shows only minor differences compared to the zero pressure gradient flow. The turbulent Schmidt number at CO2 injection shows large variations. Just downstream of the injection it attains low values, whereas high values are detected in the upper mixing zone of the cooling flow and the freestream at each film cooling row.

Large-Eddy Simulation of Supersonic Film Cooling at Finite Pressure Gradients

Large-eddy simulations are performed to analyze film cooling in supersonic combustion ramjets (Scramjets). The transonic film cooling flow is injected through a slot parallel to a Ma=2.44 main stream with a fully turbulent boundary layer. The injection Mach number is Ma i=1.2 and adiabatic wall conditions are imposed. The cooling effectiveness is investigated for adverse and favorable pressure gradients which are imposed onto the potential core region right downstream of the slot. The numerical results are in good agreement with the measured adiabatic cooling effectiveness. The turbulent mixing process of the injected cooling flow shows high turbulence levels just downstream of the lip and slowly increasing turbulence levels in the cooling flow. At a favorable pressure gradient, the adiabatic film effectiveness downstream of the potential core region is significantly increased by approximately 50% compared to the film cooling flow without a pressure gradient, whereas the adverse pressure gradient leads to a reduction of adiabatic film effectiveness by 30%.

Large-Eddy Simulation of Supersonic Film Cooling at Laminar and Turbulent Injection

To analyze lm cooling in supersonic combustion ramjets (Scramjets) Large-eddy simulations are performed. The lm cooling ow is injected into a fully turbulent boundary layer through a slot parallel to a Ma = 2:44 main stream. The injection Mach numbers are Mai = 1:8 and Mai = 2:2 at laminar and turbulent injection conditions, respectively. The blowing rate M and the total temperature ratio TR are kept constant at both injection conditions. For these injection conditions, the adiabatic cooling e ectiveness is investigated. The numerical results for the laminar injection are in good agreement with the measured adiabatic cooling e ectiveness. The turbulent injection condition leads to an increased mixing between the cooling ow and the main stream which results in an approx. 7 % reduction of adiabatic cooling e ectiveness values compared to the laminar injection condition.

Large-Eddy Simulation of Supersonic Film Cooling

Large-eddy simulations are performed to analyze lm cooling in supersonic combustion ramjets (Scramjets). The transonic lm cooling ow is injected through a slot parallel to a Ma = 2:44 main stream with a fully turbulent boundary layer. The injection Mach number is Mai = 1:2 and adiabatic wall conditions are imposed. The cooling e ectiveness is investigated for adverse and favourable pressure gradients which are imposed onto the potential core region right downstream of the slot. The numerical results are in good agreement with the measured adiabatic cooling e ectiveness. The turbulent mixing process of the injected cooling ow shows high turbulence levels just downstream of the lip and slowly increasing turbulence levels in the cooling ow. At a favorable pressure gradient, the adiabatic lm e ectiveness downstream of the potential core region is signi cantly increased by approximately 50 % compared to the lm cooling ow without a pressure gradient, whereas the adverse pressure gradient leads to a reduction of adiabatic lm e ectiveness by 30 %.

Large-Eddy Simulation of Relaminarization in Supersonic Flow

The interaction of a shock wave with an expanding supersonic turbulent boundary layer at a freestream Mach number Ma = 1:76 in a supersonic combustion ramjet inlet is analyzed using large-eddy simulations. In this context, the phenomena of relaminarization and expansion followed by compression are considered. The results are compared to a Ma = 3 computation at the same Reynolds number and an expansion angle of = 12 . Considering the relaminarization issue a skin-friction coecient reduction downstream of the expansion corner of 25 % is obtained at Ma = 1:76 and a reduction of 50 % at Ma = 3:0. Signicant Reynolds stress component reductions occur at both cases and the near wall anisotropy tends towards the one-component limit. Semi-log velocity proles reveal the formation of a large laminar-like sublayer downstream of the expansion which at the Ma = 3 case is four times longer than at the Ma = 1:76 case. The analysis of the shock wave interaction downstream of the expansion corner at Ma = 1:76 shows an 11-fold increase in the wall-normal Reynolds stress component and the tendency of the near-wall anisotropy towards the two-component limit.

Large-Eddy Simulation of Supersonic Film Cooling at Incident Shock-Wave Interaction

The impact of shock waves on supersonic cooling films is studied using large-eddy simulations (LES). A laminar cooling film is injected through a slot at a Mach number M​a i =1.8 into a fully turbulent boundary layer at a freestream Mach number M​a ∞ =2.44. The cooling film is disturbed by oblique shock waves at deflection angles of 5∘ and 8∘ at two downstream positions of the slot. At shock impingement close to the slot, i.e., within the potential-core region, at a flow deflection of 5∘, a cooling effectiveness decrease of 33% occurs downstream of the separation bubble compared to a configuration without shock impingement. If the same shock impinges further downstream upon the boundary-layer region, the decrease is only 17%. The stronger 8∘ shock wave at the further downstream impingement position leads to a maximum decrease of 33%. The current report presents a concise version of Konopka etal. (4th European conference for aerospace sciences, 2011).