Abstract The Honeywell Uncertified Research Engine (HURE), a research version of a turbofan engine that never entered production, was tested in the NASA Propulsion System Laboratory (PSL), an altitude test facility at the NASA Glenn Research Center. The PSL is a facility that is equipped with water spray bars capable of producing an ice cloud consisting of ice particles, having a controlled particle diameter and concentration in the air flow. To develop the test matrix of the HURE, numerical analysis of flow and ice particle thermodynamics was performed on the compression system of the turbofan engine to predict operating conditions that could potentially result in a risk of ice accretion due to ice crystal ingestion. The goal of the test matrix was to provide operating conditions such that ice would accrete in either the fan-stator through the inlet guide vane region of the compression system or within the first stator of the high pressure compressor. The predictive analyses were performed with the mean line compressor flow modeling code (COMDES-MELT) which includes an ice particle model. The HURE engine was tested in PSL with the ice cloud over the range of operating conditions of altitude, ambient temperature, simulated flight Mach number, and fan speed with guidance from the analytical predictions. The engine was fitted with video cameras at strategic locations within the engine compression system flow path where ice was predicted to accrete, in order to visually confirm ice accretion when it occurred. In addition, traditional compressor instrumentation such as total pressure and temperature probes, static pressure taps, and metal temperature thermocouples were installed in targeted areas where the risk of ice accretion was expected. The current research focuses on the analysis of the data that was obtained after testing the HURE engine in PSL with ice crystal ingestion. The computational method (COMDES-MELT) was enhanced by computing key parameters through the fan-stator at multiple span wise locations, in order to increase the fidelity with the current mean-line method. The Icing Wedge static wet bulb temperature thresholds were applicable for determining the risk of ice accretion in the fan-stator, which is thought to be an adiabatic region. At some operating conditions near the splitter-lip region, other sources of heat (non-adiabatic walls) were suspected to be the cause of accretion, and the Icing Wedge was not applicable to predict accretion at that location. A simple order-of-magnitude heat transfer model was implemented into the COMDES-MELT code to estimate the wall temperature minimum and maximum thresholds that support ice accretion, as observed by video confirmation. The results from this model spanned the range of wall temperatures measured on a previous engine that experienced ice accretion at certain operating conditions. The goal of this study is to show that the computational process developed on earlier engine icing tests can be used to provide an icing risk assessment in adiabatic regions for other engines.

Abstract This paper studies the influence of the high-temperature solar preheating system and chemical recuperation system on the overall performance of a MW-level three-shaft gas turbine. In order to obtain accurate simulation results, this paper adopts an integrated simulation method, where the simulation models of the solar receiver and gas turbine are established on the Matlab/Simulink platform, the simulation model of the chemical recuperation system is established on the Aspen Plus platform, and the interface between Matlab and Aspen Plus is developed for mutual communication. With the component maps, the Newton-Raphson iteration method is used to solve the thermodynamic performance of the gas turbine. Based on the above simulation models and solution method, the effects of the mean receiver wall temperature ( T s ) and steam-to-methane ratios ( R ) on the overall performance of the gas turbine are studied in this paper. From the perspective of safe operation, the influence of T s and R on the net output power of the gas turbine is first calculated. The results show that in the safe operating range, chemical recuperation can significantly increase the net output power of the gas turbine, while the net output power increases first and then decreases with the increase of T s . When the output power is constant, increasing T s or R both can effectively reduce fuel consumption, and increasing R can also improve the heat recovery efficiency and the overall thermal efficiency. However, the heat losses of the receiver increase as T s increasing, meanwhile, the heat recovery efficiency and the overall efficiency also decrease gradually. Therefore, in actual operation, various factors should be taken into account in order to reduce fuel consumption and improve the overall thermal efficiency.

Abstract The supersonic outlet conditions from a rotating detonation combustor exhibit fluctuations in temperature and pressure that exceed 200% of their mean level. Such unsteady conditions will induce a large convective heat loading onto a downstream supersonic turbine. Hence, the precise evaluation of the thermal load to the vane and rotor is essential to the design of adequate cooling strategies. In this paper, a numerical framework is proposed to compute the convective heat transfer on two types of supersonic turbines: axial and radial outflow. The fluctuations imposed at the turbine inlet were obtained from a nozzle coupled to a rotating detonation combustor. Both radial and axial turbines were designed and subsequently analyzed with full stage unsteady simulations using an Unsteady Reynolds Averaged Navier–Stokes solver. The inlet boundary conditions to the turbine are based on CFD results from a rotating detonation combustor. The unsteady adiabatic convective heat transfer coefficient was obtained from two simulations performed at a fixed homogeneous wall temperature. The heat flux variation in span-wise and stream-wise direction is analyzed in detail. Budgeting of the unsteady heat flux mechanism was performed to identify the driving contributor of the heat transfer within the turbine and finally both designs are compared.

Abstract The process of ignition in aero-engines raises many practical issues that need to be faced during the design process. Recent experiments and simulations have provided detailed insights on ignition in single-injector configurations and on the light-round sequence in annular combustors. It was shown that Large Eddy Simulation (LES) was able to reliably reproduce the physical phenomena involved in the ignition of both perfectly premixed and liquid spray flames. The present study aims at further extending the knowledge on flame propagation during the ignition of annular multiple injector combustors by focusing the attention on the effects of heat losses, which have not been accounted for in numerical calculations before. This problem is examined by developing Large Eddy Simulations of the light-round process with a fixed temperature at the solid boundaries. Calculations are carried out for a laboratory-scale annular system. Results are compared in terms of flame shape and light-round duration with available experiments and with an adiabatic LES serving as a reference. Wall heat losses lead to a significant reduction in the flame propagation velocity as observed experimentally. However, the LES underestimates this effect and leads to a globally shorter light-round. To better understand this discrepancy, the study focuses then on the analysis of the near wall region where the velocity and temperature boundary layers must be carefully described. An a-priori analysis underlines the shortcomings associated to the chosen wall law by considering a more advanced wall model that fully accounts for variable thermophysical properties and for the unsteadiness of the boundary layer.

Abstract This paper presents a numerical investigation of natural gas combustion in an OPRA OP16 tubular DLE combustor operating in premixed mode. The aim of this study is to compare different CFD models for the accurate prediction of emissions. The numerical results are validated against experimental results obtained in an atmospheric combustion test rig. A CFD workflow using AVL FIRE software has been devised enabling accurate CO and NO x emission calculations. Different turbulence models have been employed for adiabatic flow simulations at nominal load: the k-ζ-f turbulence model, PANS and LES-CSM model. Two detailed reaction mechanisms have been used to model combustion chemistry and thermodynamic properties: the DRM19 mechanism and the GRI 3.0 reaction mechanism. In the adiabatic simulations both the k-ζ-f and LES-CSM turbulence models perform well in combination with the GRI 3.0 reaction mechanism. In the non-adiabatic simulations heat exchange between combustor liner and cooling air is taken into account by using a wall temperature profile derived by scaling the experimentally obtained temperature profile such that the requirement of conservation of energy is satisfied. Reactive flow simulations at nominal load using the k-ζ-f turbulence model and GRI 3.0 reaction mechanism resulted in accurate predictions for time-averaged combustor pressure drop, CO and NO emissions: the predicted CO value deviates less than 20% from its experimentally obtained value, for NO the deviation is less than 10%. Inclusion of heat exchange between combustor and cooling air liner reduces this deviation to less than 10% for CO and less than 2% for NO. At a load equal to 106% of its nominal value, non-adiabatic simulations result in less satisfactory CO and NO predictions: NO is over predicted by approx. 20% while CO is overpredicted by nearly a factor of four. The latter is caused by the fact that experimentally obtained CO emissions reduce significantly when the load is increased from 100% to 106%, while numerically obtained CO emissions just change by a few ppm.

Abstract Time-accurate 3-D CFD simulations based on the SST-SAS turbulence model were performed to study the effects of heat load on the unsteady flow and heat transfer in a cooling duct with a staggered array of short pin fins. For this duct, the static pressure at its exit is maintained at 25 bars, and the cooling air that enters has a temperature of 673 K with a flow rate that produces a Reynolds number of 25,000. To examine the effects of heat load, the following isothermal wall temperatures were studied: 678 K, 873 K, 1073 K, and 1,273 K, which give rises to heat loads that range from 15 kW/m 2 to 1.5 MW/m 2 . Results obtained show high heat loads to cause considerable changes in the temperature of the cooling flow along the duct, which causes significant changes in density and velocity as well as viscosity and thermal conductivity. These changes along the duct were found to affect the locations where unsteady flow separation take place around the pin fins, the magnitude of the vorticity shed in the wakes, and the shedding Strouhal number. These unsteady flow mechanisms in turn strongly affect the nature of the surface heat transfer. A correlation formula for the heat transfer, which accounts for the effects of heat loads, was developed.

Abstract The high-pressure (HP) turbine and compressor blades are subjected to severe aero-thermal loads either caused by high inlet temperature resulting from the combustion chamber or by the shock wave at the blade surface. In this paper, a partitioned conjugate heat transfer (CHT) approach is used to assess 3D cases with high temperature gradient. On one hand, the high temperature resulting from the combustion chamber generates high azimuthal and radial non uniformities on the nozzle guide vane (NGV), known as hot streak. Since the engine efficiency is directly related to the turbine inlet temperature, manufacturers are seeking to improve thermal barrier coatings (TBC) to allow higher temperature. On the other hand, compressor rotor blades are also subjected to high thermal loads caused by the shock waves. Since the shock wave substantially contributes to the heat transfer, an accurate prediction of the wall temperature is thus required in order to properly estimate the rotor efficiency. However, the thermal load gradient caused by the hot streak, the use of thermal barrier coating or the shock wave have an impact on the overall computation stability. This paper assess the stability for a Dirichlet-Robin interface condition. Indeed an optimal relaxation parameter value is given for the Robin condition to provide a stable CHT computation. The impact on stability of the transient coupling time-step, the coupling frequency and the thermal conductivity is also assessed and the new transient coupling time-step defined allows to correctly stabilize the coupled simulation.

Abstract Planar Laser Induced Fluorescence (PLIF) has been demonstrated to investigate a round jet impinging on a flat surface. Detailed thermal field distributions have been obtained near the flat target surface to characterize the wall jet development ensuing from the stagnation point. While PLIF has been demonstrated for combustion applications to measure concentration gradients within a mixture, its application for temperature field measurements is less established. Therefore, the technique was applied to a simple, cylindrical impinging jet. The jet Reynolds number varied with Re jet = 5,000–15,000 while the jet–to–target surface spacing varied from H / D = 4–10. The cooling jet (T jet ∼ 300 K) impinged on a flat, heated surface. The PLIF technique was able to capture the free jet structure and jet development along the target surface. With a short impingement length (H / D = 4), the potential core of the jet strikes the target surface. The thermal gradients captured during the experiments demonstrate the fully turbulent nature of the impinging jet with H / D = 10. The thermal boundary development along the target surface is clearly captured using this fluorescence method. The near wall temperature gradients acquired with the PLIF method have been used to calculate heat transfer coefficients on the heated surface, and these values compare favorably to those measured using a well-established steady state, heat transfer method. The PLIF technique has been demonstrated for this fundamental impingement setup, and it has proven to be applicable to more complex heat transfer and cooling applications.

Abstract Aerodynamic cooling of hot components/surfaces such as those encountered in gas turbine engines is needed to avoid premature failure of parts due to thermo-mechanical stresses. An effective way of achieving this cooling is through the exchange of heat via effusion/film cooling holes on the hot surfaces. The gases absorb heat as they flow through the cooling holes and also by forming a protective layer of relatively cool gases near the hot surface. Modeling these processes allow for durable design of components and computational simulations offer a complementary way to design new parts or enable performance assessment of the existing parts at new operating conditions. However, in order to perform numerical simulations of heat transfer through effusion holes, the heat conduction through the solid liner and the convection from the gas phase must be coupled considering all the relevant length and time scales. The time scale separation between the solid and the gas phase makes this prohibitively expensive for large scale computations. In applications involving hundreds of effusion holes, resolving the geometry of each effusion hole along with the primary flow (with typically larger length scales) is very challenging. In the current work, we overcome the difficulties associated with the resolution of cooling holes by employing a local source method (Andreini et al, J. Eng. Gas Turb. Power, 2014) to model the heat transfer to the walls. This method is assessed in a canonical configuration based on experiments performed by Gustafsson (Gustaffson, Ph.D. Thesis, 2001). Large Eddy Simulations (LES) coupled with conjugate heat transfer (CHT) models are used in this study. Simulations that resolve the flow passages explicitly using mesh both in the fluid and the solid domains, were used to validate the fidelity of grid resolution, turbulence models and other simulation parameters in predicting velocity fields and wall temperature data. Although, resolving all the effusions passages provides the most accurate results, it is not practical in real applications. Hence, a local source model is employed to model the heat transfer that happens in the cooling-hole passages. In this method, the effusion passage is not resolved (using a mesh) and the mass transfer across the cooling hole passage is prescribed as an injection-extraction boundary condition. The heat transfer at the fluid-solid interface of the cooling-hole passage is also modeled based on Nusselt number correlations available in the literature. This modeling procedure enables simulations with flexible mesh topologies that can be generated at a relatively low cost in comparison to the fully resolved mesh configurations. The local source method is assessed and validated using the available experimental data. The results show that the meshes which resolve the penetration depth in the solid and which conform at the solid-gas interface provide better prediction of the wall temperatures.

Abstract Previous researches on latticework were focused on the convective heat transfer performance on pressure and suction sides of a blade model. Besides, it has an effect on leading edge by impingement. Thus, the present study provides heat transfer enhancement and pressure loss of jet impingement of a latticework on side wall in a wide channel (AR = 4). Two latticework configurations with impingement effects are employed in this study. Three kinds of sub-channel models are used in this experiment, which is according to different cooling designs. The angle of the rib is 45° and the numbers of subchannel are 4, 6 and 8, respectively. Reynolds number range is from 10000 to 30000 with an increment of 10000. The wall temperature is obtained by using wide band liquid crystal technique, and then the heat transfer coefficients on the target surface of the channel are achieved. Pressure drop of the latticework channel is also measured by pressure taps. The result shows that these two latticework models have different flow and heat transfer characteristics. The Nusselt number distribution is not similar to that of traditional jet array impingement. The range of Nusselt number enhancement is 2.3 to 6.4 compared to that of a smooth convective channel (the Nusselt number is based on the channel hydraulic diameter). The jet-to-target distance could reduce the overall averaged heat transfer on side wall. But it could also lead to a high Nu region. To the normal lattice model, more sub-channels there are, more pressure loss it has. To the novel lattice model, sub-channel number cannot affect the pressure loss, but the jet-to-target distance could affect the friction factor obviously.

Abstract In the present paper, lattice geometries have been developed and tested for application in internal cooling of gas turbine blades. These geometries are suitable for embedding within a near surface or double wall cooling configuration. Two types of unit cells with variation in number of ligaments and porosity, were used to generate two lattice configurations. The first type of unit cell consisted of six ligaments of 0.5 mm diameter joined at a common vertex situated at the middle. As such, three ligaments were located on each side of the common middle vertex. In addition, the top half of the unit cell was rotated 60 degrees compared to the bottom half of the unit cell. The second type of unit cell was derived from the first type but with four mutually perpendicular ligaments added in the middle plane. Two lattices, referred to as L1 and L2, were obtained by repeating the type 1 and type 2 unit cells, respectively, in both streamwise and spanwise directions. The test coupons consisted of these lattice structures embedded inside a channel of 2.54 mm height, 38.07 mm width and 38.1 mm in length. The coupons were fabricated using Inconel 718 powder through selective laser sintering (SLS) process. The heat transfer and pressure drop performance was evaluated using steady state tests with constant wall temperature boundary condition and for channel Reynolds number ranging from 2,800 to 15,000. In addition, steady state numerical conjugate heat transfer simulations were conducted to obtain a detailed insight into the prevalent flow field and temperature distribution. Experimental results showed a heat transfer enhancement of upto 3 times at the highest Reynolds number compared to a smooth channel. Both the heat transfer and pressure drop increased with a decrease in the porosity from L1 to L2. The numerical results revealed formation of prominent vortical structures in the inter-unit cell spaces. These vortical structures were located in the upper half of the channel causing an increased heat transfer on the top end wall. As a result, these lattice structures provided an augmented heat transfer with a potential for favorable redistribution of the coolant.

Abstract The design of heat exchangers for use in a supercritical carbon dioxide (sCO 2 ) Brayton cycle power converter must provide for acceptable performance for duty cycle events encompassing anticipated transients and postulated accidents. This paper presents the results of a comprehensive analysis of thermal transients for sCO 2 cycle heat exchangers, with emphasis on the sodium-to-CO 2 heat addition heat exchanger. A range of transients, from normal operation to severe accidents, were simulated with the coupled PDC and SAS4A/SASSYS-1 system level dynamic analysis computer codes. For each transient, the calculated change in the heat exchanger wall temperature is determined as a measure of the thermal loading.

Abstract The cooling ratio on a dynamically forced 7×7 impingement jet array is studied experimentally. The current study is focused on determining the influence of a phase-shift between every row of nozzles as well as the impact of a duty cycle variation on the cooling ratio. Both parameters are studied in dependency of the impingement distance (H/D = 2, 3, 5), the (nozzle-)Reynolds-number (Re D = 3200, 5200, 7200) and the excitation frequency (f = 0 Hz – 1000 Hz). For every set of parameters, the phase-shift between every row of nozzles is varied between Φ = 0% – 90% while the variation of the duty-cycle is performed between DC = 35% – 65%. During all investigations, the dimensionless distance between adjacent nozzles is fixed at x/D = y/D = 5 and liquid crystal thermography is used to acquire the wall temperatures, which are further processed to calculate the local Nusselt numbers. Generally, the implementation of an excitation frequency allows a case depending increase in cooling ratio of up to 52%. Further implementation of a phase-shift yields an additional frequency-depending improvement of the cooling ratio. In case of duty cycle variation, the best case revealed an additional 19% improvement in cooling ratio.

The study of boundary layer transition plays a fundamental role in the field of turbomachinery owing to its strong influence on skin friction and heat transfer. The understanding of the laminar to turbulent transition can help designers to improve the aerodynamic and thermodynamic performances both of the components and of the whole machine. Turbulent transition models are nowadays commonly used tools in both research and design practice. In the context of high-pressure turbines design, it is then of particular interest to understand if such models are able to predict the effect of temperature on bypass transition and, in case of positive answer, the reasons of their behaviour. This becomes even more interesting as the effect of the flow aero-thermal coupling becomes prominent in the analysis of such phenomena, as this effect is typically not accountedfor in the validation of turbulence models. Two state-of-the-art transition models are examined in the present contribution: the γ–Re θ model developed by Langtry and Menter [1] and the k–k l –ω model by Walters and Cokljat [2]. The two models have been chosen also as they use two radically different approaches to describe the transition process: an empirical, correlation-based one for the former model opposed to a phenomenological, based on local transport, for the latter. To isolate the effects of the temperature ratio on the transition, the simulations have been performed keeping the same values of Reynolds and Mach numbers and changing the value of the wall to free stream Temperature Ratio (TR). The results of the two transition models have been compared between them as well as with experimental results obtained as part of a parallel effort. The results show that both models are sensitive to TR and can have qualitative agreement with the observations from experimental data. Most importantly the present results show how a transition modelling based on local transport, rather than empirical correlations should be favoured.

A smooth two-pass internal gas turbine cooling channel is numerically investigated. Transient conjugated non-rotating and rotating URANS simulations are executed. The transient thermochromic liquid crystal (TLC) approach is supported with these simulations as the temporally changing rotational buoyancy effects can be examined. The Reynolds number is 25,000, the rotation number is 0.24 and the initial buoyancy number is 0.63 (according to an inlet-to-wall density ratio of 0.23). As heat is transferred, the temperatures and heat fluxes change with increasing time, and so do the local buoyancy effects. The computational results are evaluated as averaged segmental values. They are compared to the experimental results from literature that have been determined for constant wall temperature experiments (various experimental runs with different constant wall temperatures). Especially in the first passage, there is a good agreement between the numerically gained results and the experimental data. The more complex flow inside the bend leads to more diverse characteristics and the second passage is only slightly effected.

With the goal to realize gas turbine engine cycle efficiency of 65%, the turbine inlet temperature is predicted to rise to 1700° C. Additive manufacturing (AM) has opened new avenue to explore complex design spaces to enhance cooling in gas turbine blades. Present paper developed a series of additively manufactured parallel cooling channels with streamwise wall jets inside, which could be suitable for double wall and near wall cooling configurations in gas turbine hot section components. The tested coupons consisted of parallel channels, each channel further divided into small chambers using several spanwise separation walls. Height of these walls was kept less than channel height, thus forming a crossover slot jet with one of the end walls. Coolant entered from one side of channel and formed streamwise wall jet while crossing through the slot over to the downstream chamber. The test coupons were additively manufactured by Selective Laser Sintering technique using Inconel 718 alloy. Both experimental and numerical study has been conducted to analyze effect of pitch and blockage ratio on heat transfer. Pitch between subsequent slots was defined in terms of slot hydraulic diameter and was varied from 5D j to 11 D j . Blockage ratio consisted of ratio of separation wall height to the channel height and varied from 0.5 to 0.75. Steady state heat transfer experiments with constant wall temperature boundary condition were performed for a channel Reynolds number range of 1800 to 5000. ANSYS CFX solver with the SST k-ω turbulence model was used to perform numerical simulation to obtain detailed understanding of existing flow field. Experimental results showed heat transfer enhancement of up to 6.5 times that of a smooth channel for the highest blockage ratio of 0.75. The blockage ratio exerted stronger effect both on pressure drop and heat transfer characteristics, as compared to pitch. The heat transfer as well as pressure drop were found to increase with a decrease in pitch and with an increase in blockage ratio. The numerical results revealed complex flow field which consisted of wall jets along with impingement, separation and recirculation zones in each chamber. It also revealed significant uniformity in heat transfer coefficient distribution on the slot side end wall. For all configurations, the gain in heat transfer was accompanied with high pressure drops. However, coupled with the high heat transfer, this design could lead to potential coolant savings.

Soot production in turbulent flames represents an urgent issue for applied systems and raises a range of fundamental challenges. Most of the literature effort on this question has been carried out on non-premixed turbulent flames, where mixing plays a dominant role. It is however interesting to see if soot production can be investigated in a premixed turbulent flame configuration, thus eliminating the role of mixing in this process. It is shown in the present article that this can be accomplished by making use of rich swirled ethylene/air flames under perfectly-premixed rich conditions established in a confined combustor operating at atmospheric pressure. Quantitative measurements of the soot volume fraction (f v ) are carried out by the Laser Induced Incandescence (LII) technique. Although the soot volume fraction levels are one or two orders of magnitude lower than those found in non-premixed flames, it is shown that detection is feasible and that this configuration may be used to analyze effects on soot production of operating parameters such as the equivalence ratio, the power level or the wall temperature. The observed trends are interpreted numerically by considering an idealized model of a rich premixed burner-stabilized stagnation flame. This configuration is here calculated with a detailed kinetic scheme in combination with a soot sectional model. This description is able to account for some of the trends observed experimentally and indicates in particular how an increase in the wall temperature may increase the soot volume fraction as it is effectively observed in the experiment. The model predicts that the soot volume fraction level increases for high equivalence ratios. This finding is in disagreement with the experimental observations on this swirled flame, which exhibits a maximum soot volume fraction for ϕ = 2.1 for all the considered flame powers. This indicates that complex interactions take place between soot, flame, turbulence and thermal environment and that the investigation of these processes will require a comprehensive turbulent simulation. This may be guided by the database developed in the present experimental effort.

Phosphor thermometry has been developed for wall temperature measurements in gas turbines and gas turbine model combustors. An array of phosphors has been examined in detail for spatially and temporally resolved surface temperature measurements. Two examples are provided, one at high pressure (8 bar) and high temperature and one at atmospheric pressure with high time resolution. To study the feasibility of this technique for full scale gas turbine applications a high momentum confined jet combustor at 8 bar was used. Successful measurements up to 1700 K on a ceramic surface are shown with good accuracy. In the same combustor, temperatures on the combustor quartz walls were measured, which can be used as boundary conditions for numerical simulations. An atmospheric swirl-stabilized flame was used to study transient temperature changes on the bluff body. For this purpose, a high-speed setup (1 kHz) was used to measure the wall temperatures at an operating condition where the flame switches between being attached (M-flame) and being lifted (V-flame) (bistable). The influence of a precessing vortex core (PVC) present during M-flame periods is identified on the bluff body tip, but not at positions further inside the nozzle.

Bearing chamber of a gas turbine engine is generally sealed by pressurized air, separating lubricant from the other zones of the engine. Heat transfer from the wall to air/oil mixture is a challenging engineering problem; predicting heat transfer rate from bearing chamber to oil is important to avoid oil coking and oil fires under high rotational speeds, pressure levels and turbine inlet temperatures. In this study, the inner wall temperature of bearing chamber which is located at the center of front engine structure was investigated numerically. The numerical study involved mainly two thermal modelling methods having two different empirical correlations was performed with finite element solver in order to calculate heat transfer on the wall. First method was based on rotational Reynolds number and Prantl number, in addition to these numbers second one, which is suggested in the literature, is based on oil related and sealing air related Reynolds number, mixture temperature and mixture mass flow. Second approach considers existence of a mixing of gaseous and liquid flow in the core flow unlike first modelling approach. The thermal model was solved by finite element solver and numerical model, assumptions were described with thermal boundary conditions. On the other hand, wall and air thermocouple readings were taken through engine test from the bearing chamber for real engine operating conditions having mainly idle, cruise and maximum power. DN number ranges from 712564 to 2742404, sealing air flow ranges from 46 to 78 g/s and oil flow ranges 22 to 40 g/s for these conditions. The calculated heat transfer coefficients were presented and discussed. The wall temperature predictions of the thermal models, and test measurements were compared. The comparison revealed that analysis results obtained with both correlations were in reasonable agreement with the test. In overall, the second approach predicted metal temperature slightly better at the front support and inner manifold wall, while first approach predicted much better at the rear support wall.

Separated boundary layers of the low pressure turbine blade suction surface under wall heat transfer and pressure gradient conditions are investigated using large eddy simulations (LES) in this paper. The study constructed a converging-diverging channel with a flat plate as the bottom wall, and the pressure distribution of the bottom wall is similar to that of a high lift low pressure turbine blade suction surface. The boundary layer was investigated under different heat transfer boundary conditions of the bottom wall (i.e., the adiabatic wall and the isothermal wall with the wall temperature being 0.8 times of the inflow temperature). The time-averaged flow parameters and the separation bubble characteristics were analyzed and discussed. The evolution of coherent structure diagrams of the boundary layer was also obtained to study the evolution process of the vortex. The results show that the cooled isothermal wall condition can significantly suppress the separation bubble and reduce the frequency of the large scale spanwise vortex roll-up. Under the two wall heat transfer conditions, the scale of the near wall small scale spanwise vortex is similar as well as the scale of the large scale spanwise wortex. The location of the vortex are also approximate under the two wall heat transfer conditions, but the position of the large scale spanwise vortex shedding from separated laminar boundary layers moves upstream under the cooled isothermal wall condition, and the transition process is more rapid than that of the adiabatic wall condition.