No abstract

Lean burn swirl stabilized combustors represent the key technology to reduce NOx emissions in modern aircraft engines. The high amount of air admitted through a lean-burn injection system is characterized by very complex flow structures such as recirculations, vortex breakdown and processing vortex core, that may deeply interact in the near wall region of the combustor liner. This interaction and its effects on the local cooling performance make the design of the cooling systems very challenging, accounting for the design and commission of new test rigs for detailed analysis. The main purpose of the present work is the characterization of the flow field and the wall heat transfer due to the interaction of a swirling flow coming out from real geometry injectors and a slot cooling system which generates film cooling in the first part of the combustor liner. The experimental setup consists of a non-reactive three sector planar rig in an open loop wind tunnel; the rig, developed within the EU project LEMCOTEC, includes three swirlers, whose scaled geometry reproduces the real geometry of an Avio Aero PERM (Partially Evaporated and Rapid Mixing) injector technology, and a simple cooling scheme made up of a slot injection, reproducing the exhaust dome cooling mass flow. Test were carried out imposing realistic combustor operating conditions, especially in terms of reduced mass flow rate and pressure drop across the swirlers. The flow field is investigated by means of PIV, while the measurement of the heat transfer coefficient is performed through Thermochromic Liquid Crystals steady state technique. PIV results show the behavior of flow field generated by the injectors, their mutual interaction and the impact of the swirled main flow on the stability of the slot film cooling. TLC measurements, reported in terms of detailed 2D heat transfer coefficient maps, highlight the impact of the swirled flow and slot film cooling on wall heat transfer.

State-of-the-art liner cooling technology for modern combustion chambers is represented by effusion cooling (or full-coverage film cooling). Effusion is a very efficient cooling strategy typically based on the use of several inclined small diameter cylindrical holes, where liner temperature is controlled by the cornbined protective effect of coolant film and heat removal through forced convection inside each hole. A CFD-based thermal analysis of such components implies a significant computational cost if the cooling holes are included in the simulations; therefore many efforts have been made to develop lower order approaches aiming at reducing the number of mesh elements. The simplest approach models the set of holes as a uniform coolant injection, but it does not allow an accurate assessment of the interaction between hot gas and coolant. Therefore higher order models have been developed, such as those based on localized mass sources in the region of hole discharge. The model presented in this paper replaces the effusion hole with a mass sink on the cold side of the plate, a mass source on the hot side, whereas convective cooling within the perforation is accounted for with a heat sink. The innovative aspect of the work is represented by the automatic calculation of the mass flow through each hole, obtained by a run time estimation of isentropic mass flow with probe points, while the discharge coefficients are calculated at run time through an in-house developed correlation. In the same manner, the heat sink is calculated from a Nusselt number correlation available in literature for short length holes. The methodology has been applied to experimental test cases of effixsion cooling plates and compared to numerical results obtained through a CFD analysis including the cooling holes, showing a good agreement. A comparison between numerical results and experimental data was peiformed on an actual combustor as well, in order to prove the feasibility of the procedure

No abstract

State-of-the-art liner cooling technology for modern combustors is represented by effusion cooling (or full-coverage film cooling). Effusion is a very efficient cooling strategy based on the use of multiperforated liners, where the metal temperature is lowered by the combined protective effect of the coolant film and heat removal through forced convection inside each hole. The aim of this experimental campaign is the evaluation of the thermal performance of multiperforated liners with geometrical and fluid-dynamic parameters ranging among typical combustor engine values. Results were obtained as the adiabatic film effectiveness following the mass transfer analogy by the use of pressure sensitive paint, while the local values of the overall effectiveness were obtained by eight thermocouples housed in as many dead holes about 2 mm below the investigated surface. Concerning the tested geometries, different porosity levels were considered: such values were obtained by both increasing the hole diameter and pattern spacing. Then the effect of the hole inclination and aspect ratio pattern shape were tested to assess the impact of typical cooling system features. Seven multiperforated planar plates, reproducing the effusion arrays of real combustor liners, were tested, imposing six blowing ratios in the range 0.5–5. Additional experiments were performed in order to explore the effect of the density ratio (DR=1;1.5) on the film effectiveness. Test samples were made of stainless steel (AISI304) in order to achieve the Biot number similitude for the overall effectiveness tests. To extend the validity of the survey a correlative analysis was performed to point out, in an indirect way, the augmentation of the hot side heat transfer coefficient due to effusion jets. Finallyv,in order to address the thermal behavior of the different geometries in the presence of gas side radiation, additional simulations were performed considering different levels of radiative heat flux.

No abstract

State-of-the-art liner cooling technology for modern combustors is represented by effusion cooling (or full-coverage film cooling). Effusion is a very efficient cooling strategy based on the use of multi-perforated liners, where metal temperature is lowered by the combined protective effect of coolant film and heat removal through forced convection inside each hole. The aim of this experimental campaign is the evaluation of the thermal performance of multi-perforated liners with geometrical and fluid-dynamic parameters ranging among typical combustor engine values. Results were obtained as adiabatic film effectiveness following the mass transfer analogy by the use of Pressure Sensitive Paint, while local values of overall effectiveness were obtained by eight thermocouples housed in as many dead holes about 2 mm below the investigated surface. Concerning the tested geometries, different porosity levels were considered: such values were obtained both increasing the hole diameter and pattern spacing. Then the effect of hole inclination and aspect ratio pattern shape were tested to assess the impact of typical cooling system features. Seven multi perforated planar plates, reproducing the effusion arrays of real combustor liners, were tested imposing 6 blowing ratios in the range 0.5–5. Test samples were made of stainless steel (AISI304) in order to achieve Biot number similitude for overall effectiveness tests. To extend the validity of the survey a correlative analysis was performed to point out, in an indirect way, the augmentation of hot side heat transfer coefficient due to effusion jets. Finally, to address the thermal behaviour of the different geometries in presence of gas side radiation, additional simulations were performed considering different levels of radiative heat flux.

This paper reports the numerical study performed during the setup of a tubular laboratory combustor developed by Avio Group. The main purpose of this study is the assessment of some numerical tools which might be employed for the design of aero-engines combustors. Adopted methodologies pertains with heat transfer and exhaust emissions issues of combustor design, and they refer to both detailed three dimensional and simplified zero/one dimensional formulations. An exhaustive discussion about the obtained result and about the adopted modelling criteria is reported in the paper. The rig has a tubular geometry that allows to study the single burner of a typical annular combustor of jet engines. Considered operating condition for the tests refers to Max Take-Off (ICAO 100%) in the cycle of a typical jet engine. Three different numerical tools were considered, representing the fundamental steps during the preliminary and detailed design of combustors. The first one is a 1D procedure capable to analyze the cooling flow network of the combustor and to predict liner wall temperature and heat loads. The second is a classical chemical reactor network code for flame temperature and exhaust emissions prediction. Such preliminary design tools are usually supported by reactive CFD computations. In this work both two dimensional and three dimensional CFD models are considered with different goals. A 2D CFD model of flame tube with prescribed mass flow rates allows to quickly test various turbulence, combustion and emissions models, while a 3D model of the complete hardware (liner, burner and cooling network), permits to predict air flow splits and wall radiative and convective heat loads. Furthermore a final 3D reactive CFD computation with radiation conjugated with the thermal conduction solution across the liner, permits to estimate the wall temperature distribution. The whole set of numerical results has pointed out an appreciable agreement among the various tools representing a valid assessment of the validity and robustness of selected design methodologies. A more significant validation of codes accuracy will be possible as soon as the scheduled experimental results will be available.