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The influence of film-hole position on internal and external heat transfer was investigated using Computational Fluid Dynamics (CFD). A simplified geometry of an integrated configuration of a ribbed channel, film hole and mainstream passage is modeled to represent a turbine internal and external cooling scheme. The proposed configurations with nine different positions of film holes are parameterized to conduct a series of CFD calculations at a target blowing ratio of 0.8, 1.1 and 1.7. Since the present study is taking a comparative approach, CFX with SST models is applied as a primary tool and the results are compared with Fluent solver for selected cases (total 36 cases). Among the proposed nine positions, the film holes located in the separated flow region of a ribbed channel showed considerable enhancement in film effectiveness with minimum reduction and potential improvement in internal heat transfer. The finding offers a design opportunity to enhance internal as well as external heat transfer. Published version

The goal of this study is to propose a rapid design approach for designing a minimum pressure loss U-bend in a turbine internal cooling passage using topology optimization. The total pressure loss in the bend region is critical, as the presence of U-bend dominates the pressure loss so that it demands increased coolant feed pressure. However, it is challenging to come up with general design rule for U-bend with minimum pressure loss. The proposed rapid design approach can be applied as a 3D U-bend geometry optimization tool. The minimization of the total pressure loss is achieved by implementing topology optimization that uses a continuous adjoint approach with steepest-descent method. An incompressible steady-state flow (low-fidelity simulation) solver is applied for the design space, which is modeled as a porous medium with variable porosity. A series of design optimization in 2D and 3D is conducted at a Reynolds number of 100,000 based on the bulk inlet velocity. 3D U-bend configurations are produced by the proposed rapid design approach in a few hours. High-fidelity computational simulations are also carried out on the optimized 2D and 3D U-bend configurations to evaluate the proposed rapid design approach. The computational studies are performed by solving the compressible, turbulent steady-state Reynolds-averaged Navier-Stokes (RANS) equations applying two turbulence models: Spalart-Allmaras and Shear Stress Transport (SST) model. The high-fidelity simulations predict a relative improvement of maximum 26.8% in total pressure drop with respect to the 3D baseline configuration. The results indicate that the proposed approach can be reliable fast design tool for optimizing U-bend geometry in incompressible flow regime.

The purpose of this investigation was to explore the effects of leakage flows from labyrinth seals in shrouded stators on the performance and detailed flow characteristics of multistage axial compressors. A three-stage axial compressor with an inlet guide vane operating under transonic conditions was considered. The three-dimensional unsteady Reynolds-averaged Navier–Stokes equation with a shear stress transport turbulence model was employed in the simulation. The transient rotor–stator model was used at the interface between the rotor and stator to consider the flow interaction and propagation from upstream to downstream. The results clearly presented that the leakage flows could reduce the compressor performance due to a disturbance in the main flow. Moreover, it was observed that the leakage, which flowed from the labyrinth seal to the main flow path, can intensify the recirculation characteristics, such as the hub-corner stall. On the other hand, the reversed leakage flows, corresponding to flow from the main flow path to the seal side, reduced the size of the hub-corner stall. In addition, the spanwise total pressure loss characteristics were altered by the size of the recirculation flows due to leakage flows. Thus, the leakage flows could affect the total pressure loss and blockage, which are important parameters in compressor design.

Predicting the exact performance of multi-stage compressors for an aircraft gas turbine is important in the design process and calculation of off-design characteristics. This study explores the quantitative effects of cold-to-hot deformation on the aerodynamic performance of a three-stage axial compressor in terms of power consumption and efficiency. The reason of difference between numerical result and performance test data is investigated. Three-dimensional computational fluid dynamics (3D CFD) was used to predict the flow characteristics. A finite-element (FE) analysis was conducted to calculate the displacement caused by rotation, thermal expansion, and pressure force. First, the fluid domain of the compressor with cold geometry was analyzed via a 3D CFD analysis, and the displacement for all coordinates was then obtained from the FE analysis. The hot geometry and mesh for the fluid domain were generated based on the cold geometry and calculated displacement. Finally, the aerodynamic performance of the compressor with hot geometry was evaluated via a 3D CFD analysis. The results for the hot geometry compressor were more consistent with the performance test data than those for the cold geometry compressor in terms of the pressure ratio and efficiency based on the mass flow rate. Moreover, the average rotor tip clearance decreased by approximately 0.5% with respect to the span height by the cold-to-hot deformation. The total pressure loss at the rotor tip region was lower for the hot geometry compressor than the cold geometry compressor because of the reduced entropy of the high vortex loss core around the rotor tips. The results showed that the power consumption was 1.16% higher for the hot geometry compressor than the cold geometry compressor at the designed pressure ratio, which was primarily caused by an increase in the mass flow rate due to the wider throat area after the cold-to-hot deformation.

An adaptation method was proposed to improve the modeling accuracy of overall and local performances of gas turbine engine. The adaptation method was divided into two steps. First, the overall performance parameters such as engine thrust, thermal efficiency, and pressure ratio were adapted by calibrating compressor maps, and second, the local performance parameters such as temperature of component intersection and shaft speed were adjusted by additional adaptation factors. An optimization technique was used to find the correlation equation of adaptation factors for compressor performance maps. The multi-island genetic algorithm (MIGA) was employed in the present optimization. The correlations of local adaptation factors were generated based on the difference between the first adapted engine model and performance test data. The proposed adaptation method applied to a low-bypass ratio turbofan engine of 12,000 lb thrust. The gas turbine engine model was generated and validated based on the performance test data in the sea-level static condition. In flight condition at 20,000 ft and 0.9 Mach number, the result of adapted engine model showed improved prediction in engine thrust (overall performance parameter) by reducing the difference from 14.5 to 3.3%. Moreover, there was further improvement in the comparison of low-pressure turbine exit temperature (local performance parameter) as the difference is reduced from 3.2 to 0.4%.

Aim of this work is to provide evidence of the effectiveness of combined use of the variable guide vanes (VGVs) and bleed air on the thermodynamic performance of aircraft engine system. This paper performed the comparative study to evaluate the overall thermal performance of an aircraft engine with optimized VGVs and bleed air, separately or simultaneously. The low-bypass ratio turbofan engine has been modeled with a 0D/1D modeling approach. The genetic algorithm is employed to find the optimum schedule of VGVs and bleed air. There are four types of design variables: (1) the inlet guide vane (IGV) angle, (2) the IGV and 1st stator vane (SV) angles, (3) bleed air mass flow rate at the exit of the axial compressor, and (4) both type 2 and type 3. The optimization is conducted with surge margin constraints of more than 10% and 15% in the axial compressor. The results show that the additional use of the bleed air increases the efficiency of the compressors. Overall, the percentage reductions of the total fuel consumption for the engine with the IGV, 1st SV and bleed air schedule is 1.63% for 15% surge margin constraints when compared with the engine with the IGV schedule.

Steady Reynolds-averaged Navier-Stokes (RANS) simulation with the mixing-plane approach is the most common procedure to obtain the performance of a centrifugal compressor in an industrial development process. However, the accurate prediction of complicated flow fields in centrifugal compressors is the most significant challenge. Some phenomena such as the impeller-diffuser flow interaction generates the unsteadiness which can affect the steady assumption. The goal of this study is to investigate the differences between the RANS and URANS simulation results in a centrifugal compressor stage. Simulations are performed at three operating points: near surge (NS), design point (DP), and near choke (NC). The results show that the RANS simulation can predict the overall performance with reasonable accuracy. However, the differences between the RANS and URANS simulation are quite significant especially in the region that the flows are highly unsteady or nearly separated. The RANS simulation is still not very accurate to predict the time-dependent quantities of the flow structure. It shows that the URANS calculations are necessary to predict the detailed flow structures and performance. The phenomena and mechanisms of the complex and highly unsteady flow in the centrifugal compressor with a vaned diffuser are presented and analyzed in detail. Published version

In this study, we perform a series of aero-thermo-mechanical analyses to predict the running-tip clearance and the effects of impeller deformation on the performance using a centrifugal compressor. During operation, the impeller deformation due to a combination of the centrifugal force, aerodynamic pressure and the thermal load results in a non-uniform tip clearance profile. For the prediction, we employ the one-way fluid-structure interaction (FSI) method using CFX 14.5 and ANSYS. The predicted running tip clearance shows a non-uniform profile over the entire flow passage. In particular, a significant reduction of the tip clearance height occurred at the leading and trailing edges of the impeller. Because of the reduction of the tip clearance, the tip leakage flow decreased by 19.4％. In addition, the polytrophic efficiency under operating conditions increased by 0.72％. These findings confirm that the prediction of the running tip clearance and its impact on compressor performance is an important area that requires further investigation.

A series of aero-thermo-mechanical analyses were carried out to predict the running tip clearance and the effects of impeller deformation on the performance using two different centrifugal compressors (blade type A and B). In operation, impeller deformation due to the combination of centrifugal force, aerodynamic pressure and thermal load results in non-uniform tip clearance profile. The results show that the maximum displacement occurs at the leading edge tip of the impeller blade but maximum stress takes place at the blade root of the impeller. A significant reduction of the tip clearance height has occurred at the leading edge and the trailing edge of the impeller. Due to the reduction of the tip clearance, the tip leakage flow has decreased by 19.4% and 16.2% in the blade type A and B, respectively. The polytropic efficiency of blade type A and B at operating condition has increased by 0.72% and 1.81%, respectively.

In this study, an analysis of steady-state and unsteady-state Conjugate Heat Transfer (CHT) of an aeronautic high pressure gas turbine was conducted, which can calculate fluid and solid domain at the same time. ANSYS CFX V16.0 was used to solve the problem. The main emphasis of this study was on three dimensional behavior of the temperature distribution in blade of the 1st stage high pressure turbine. The Conjugate heat transfer approach in this study is validated with 1983 NASA internally cooled C3X experimental data. Results from the unsteady state were compared to the results of steady state calculations. This paper focuses on the need of unsteady conjugate heat transfer analysis of the rotor blade in cooling design process, with the prediction of the thermal environment around the rotor blade and heat conduction in the rotor blade as it is quite important to carry out the thermal load analysis and life assessment of rotor blades.

Diverse cross-corrugated surface geometries were considered to estimate the sensitivity of four variants of k-e turbulence models (Low Reynolds, standard, RNG and realizable models). The cross-corrugated surfaces considered in this study are a conventional sinusoidal shape and two different asymmetric shapes. The numerical simulations using the steady incompressible Reynolds-averaged Navier- Stokes (RANS) equations were carried out to obtain the steady solutions of the flow and thermal fields in the unitary cell of the heat exchanger matrix. In addition, the experimental test for the measurement of local convective heat transfer coefficients on the heat transfer surfaces was performed by means of the Transient liquid crystal (TLC) technique in order to compare the numerical results with the measured data. The features on detailed flow structure and corresponding heat transfer in the unitary cell of the matrix type heat exchanger are compared and analyzed against four different turbulence models considered in this study.

Modern aircraft is facing the increase of power demands and thermal challenges. In accordance with the application of more electric technology and advanced mission requirement, aircraft system requires increase of power generation and it cause increase of internal heat generation. Simultaneously, restrictions have significantly been imposed to the thermal management system. Modern aircraft must maintain low radar observability and infra-red signature. In addition, new composite aircraft skins have reduced the amount of heat that can be rejected to the environment. The combination of these characteristics has increased the challenges faced by thermal management. In order to mitigate the thermal challenges, the concept of system level thermal management should be applied and new modeling and simulation tools need to be developed. To develop and utilize system level thermal management technology, three key points are considered. Firstly, the performance changes of subsystems and components must be assessed at an integrated thermal system. It is because that each subsystem and component interacts with other subsystems or components and it can directly effects on overall system performance. Secondly, system level thermal management requirements and solutions must be evaluated early in conceptual design process as vehicle and propulsion system configuration decisions are being made. Finally, new component level thermal management technologies must focus on reducing heat generation and increasing the availability of heat sinks.Key Words : System Level Thermal Management, Fuel Thermal Management System, Adaptive Power Thermal Management System, Heat Load, Modeling and Simulation

The present work reports the investigation on the cooling performance of rotor ventilation duct of turbine generator. An investigation using experimental and analytical approaches was carried out to understand the flow and heat transfer characteristics of a single rotor ventilation duct. For the investigation, the stationary experiments are carried out for the three different conditions of (1) no cross flow (2) inlet cross flow (3) inlet and outlet cross flow followed by CFD analysis for each case. Transient heat transfer experiment using thermochromic liquid crystal (TLC) technique was carried out for measuring heat transfer characteristics. Furthermore, pressure losses are measured. A series of CFD analysis is carried out for validating modeling approach. The CFD results show good agreement with experimental data. Since the experiments were carried out on the stationary condition, the effect of rotation is examined using validated CFD model. Reynolds number is varied in the range of 13100 ~ 21000.

Turbine generators operate with complex cooling systems due to the challenge in controlling the peak temperature of the stator bar caused by Ohm loss, which is unavoidable. Therefore, it is important to characterize and quantify the thermal performance of the cooling system. The focus of the present research is to investigate the heat transfer and pressure loss characteristics of a typical cooling system, the so-called stator ventilation duct. A real scale model was built at its operating conditions for the present study. The direction of cooling air was varied to consider its operation condition, so that there are: (1) outward flow, and (2) inward flow cases. In addition, the effect of (3) cross flow (inward with cross flow case) was also studied. The transient heat transfer method using thermochromic liquid crystals is implemented to measure full surface heat transfer distribution. A series of computational fluid dynamics (CFD) analyses were also conducted to support the observation from the experiment. For the outward flow case, the results suggest that the average Nusselt numbers of the 2nd and 3rd ducts are at maximum 100% and 30% higher, respectively, than the inward flow case. The trend was similar with the effect of cross flow. The CFD results were in good agreement with the experimental data.

In this paper, a simulation approach using a data-driven model to predict the performance of a gas turbine for aircraft engines during transient operations is proposed. The low bypass ratio and mixed-flow turbofan engine is considered in the simulation. The input parameters for the training of the data-driven model are fuel mass flow rate, altitude, flight Mach number, and the required power due to the moments of inertia of the rotating parts. The output parameters in the data-driven model are engine net thrust, low-pressure shaft rotating speed, pressure and temperature at each station, propulsion efficiency, efficiency of energy conversion and the overall efficiency. The data-driven model is trained using the data set obtained from a validated first principle model. In the simulation, using the data-driven model, the final engine performance is calculated using an iterative calculation for converging the power balance equation that considers the required power due to the moments of inertia at each time step. The transient performance simulation is tested during a throttle frequency sweep maneuver. In a comparison between the first principle model and the proposed simulation approach, the R-squared values of the output parameters are higher than 0.98, except for the efficiency of energy conversion and the overall efficiency, which register 0.9715 and 0.9183, respectively. It has been confirmed that the proposed simulation approach using the data-driven model can be applied to the transient simulation.