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The effect of the swirl clocking on three-dimensional nozzle guide vane (NGV) is investigated using computational fluid dynamics. The research reports the loss characteristics of leaned and swept NGVs and the influence of swirl clocking. The three-dimensional NGVs are built by stacking the same 2D profile along different linear axes, characterized by different angles with respect to the normal or radial direction: ε = −12° ~ +12° for the leaned and γ = −5° ~ +10° for the swept airfoils. A total of 40 models are analyzed to study the effects of lean and sweep on aerodynamic performance. To investigate the influence of swirl clocking, the analysis cases include the center of the swirl that was positioned at the leading edge as well as the middle of the passage. The prediction results show that the relationship of the changes in mass flow rate and throat area are not monotonic. Further observation confirms the redistribution of loading and flow angle under different lean and sweep angles, thus, three-dimensional design is a key influencing factor on aerodynamic performance. In the presence of swirl clocking, NGV performance is changed significantly and the findings offer new insight and opportunities to improve three-dimensional NGV airfoil design.

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.

Turbine generator operates with complex cooling system 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 characterise and quantifies the thermal performance of the cooling system. The focus of the present research is to investigate the heat transfer and pressure loss characteristics of typical cooling system, so-called stator ventilation duct. A real scale model was built at its operating conditions for the present study. The direction of cooling air is 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) is 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 analysis is also conducted to support the observation from the experiment. For the inward flow case, the results suggest that the average Nusselt number of the 2nd duct is about 30% higher than the 3rd duct. The trend is similar with the effect of cross flow. The CFD results are in good agreement with the experimental data.

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.

The present study focuses on evaluating the thermal performance of the realistic leading edge cooling passage of a turbine blade, experimentally. The detailed heat transfer distributions and pressure loss of the cooling passage are measured with and without rib turbulators for direct comparison. The experimental model is representing the realistic leading edge cooling passage, which has a converging duct with triangular cross sectional shape. Therefore, the aspect ratio and the cross sectional area of the triangular passage vary significantly along the coolant flow direction. A Perspex test model is designed at 1.3 times large scale to meet Reynolds number similarity. The test model is installed with 45-degree rib turbulators at constant space-to-height ratio of P/e = 10. The experiments are carried out at the Reynolds numbers of 20,000, 40,000 and 80,000. The transient heat transfer test technique using thermochromic liquid crystals (TLCs) was applied to measure the detailed surface heat transfer distributions. The local mixed bulk temperatures (Tmb) are defined using the air thermocouples positioned in the centerline along the cooling passage. The static pressures are measured from the pressure taps machined as a part of the test model so that friction factors of the ribbed cooling passage are also evaluated. The friction factors are influenced by the characteristics of 3 dimensional converging duct of the experimental model. The full surface heat transfer coefficients, friction factors, and overall thermal performance of the cooling passage are reported for the comparison with the other cooling configurations.

In general, thermal residual stress is generated in casting products as a result of non-uniform cooling, phase transformation from liquid to solid states, and thermo-mechanical constraints. Thus, understanding accurate stress distributions is important in designing a casting part to predict its real structural stability. However, the residual stress that occurs in the casting process is often not considered in the mechanical design process to reduce engineering efforts and time. We conducted a structural analysis that compares two cases with and without considering casting residual stress in large-scale sand castings and obtained significantly different results. We introduced an interfacing algorithm (casting-structural interface) to map the casting residual stress with structural stress and combined the results of finite different analysis for casting and finite element analysis for structural design. This work represents our preliminary studies on improving the design accuracy of a large-scale casting part and shows the importance of considering the casting residual stress. A difference of about 41% in maximal principle stress is obtained with and without considering the casting residual stress.

The pressure drop predicted by the well known Zukauskas correlation, which is used in the design of a heat exchanger with a tube bundle, was compared with measured values in order to clarify the effect of acoustic resonance on the pressure drop in a heat exchanger. A staggered tube array of 1.5 mm diameter tubes was tested over a wide range of flow velocities. Additional tests were carried out with higher natural acoustic frequencies, achieved using anti-resonance baffles, to allow the detailed pressure drop behaviour during the lock-in and unlocked phases to be assessed. The actual pressure drop was significantly higher that the pressure drop predicted using the Zukauskas correlation during the lock-in phase, and the difference between the measured and the predicted values varied up to about 40% of the measured value.

This paper proposes a full engine cycle analysis to derive the schedule of variable guide vanes (VGVs) in a multi-stage axial compressor for improving the performance of a turbofan engine with required surge margin. Most researchers assumed the operating line of a compressor in process of scheduling of VGVs. However, it has been known that the performance of a compressor is influenced by the angle of VGVs. As a result, the performance variation of a compressor can affect an engine. Therefore, a full engine cycle analysis is conducted to consider the effect of VGVs on a turbofan engine. The VGVs are applied to the front stages of the high pressure compressor in the low-bypass ratio turbofan engine. The compressor of the front stages is transonic three-stage axial flow compressor. The commercial engine simulation program, NPSS TM is employed to analyze the on- and off-design performance of the turbofan engine. 1D meanline analysis is used for performance prediction of a multi-stage axial flow compressor with VGVs. The engine simulation program is coupled with the compressor performance prediction tool to consider the engine performance variation with application of the VGVs. The proposed scheduling algorithm of the inlet guide vane (IGV) and stator vanes (SV) is adjusting the vanes angle to have the minimum loss incidence angle at all rotor blades and to satisfy the required surge margin at off-design condition. The result shows that the proposed algorithm can obtain the scheduling of the IGV, 1st and 2nd SVs angle with stable operation and improved specific fuel consumption of the engine.

To enhance the thermal efficiency of aero-engines, many works on numerical and experimental evaluation of cross corrugated heat exchangers for applications such as intercoolers and recuperators have been reported. However, numerical analyzes of unit cells or of corrugated flow passages in cross corrugated heat exchangers cannot be compared directly with experimental analyzes performed only at inlets and outlets. This paper presents results of the direct measurement of static pressure distribution in a corrugated flow passage: measurement was performed by embedding pressure tapping holes connected with microchannels into the corrugated plate, which had a thickness of 200 μm. Four corrugated plates were stacked and brazed to manufacture a real scale cross corrugated heat exchanger. The static pressure distribution was successfully measured inside the compact and complex heat exchanger. The friction factors for different Reynolds numbers were correlated for hydrodynamic performance evaluation. Finally, the friction factor was discussed with a consideration of the accuracy of the manufacturing of the cross corrugated heat exchanger.

The U-bend which turns flow through 180° is encountered in many applications in mechanical and aerospace engineering systems. One important example occurs in modern turbine blade cooling systems, where the internal cooling passages is threaded radially outwards and inwards to form a multi-pass arrangement where straight passages are connected with U-bends. The main purpose of the present investigation was to focus on finding a 3D U-bend configuration with minimum pressure loss using the 3D CAD-based surface parameterisation method. The design of experiment technique and surrogate design space model were successfully applied by the authors, as opposed to direct numerical optimisation, to reduce the computational cost. A standard Reynolds-averaged Navier–Stokes (RANS) computational fluid dynamics method with the Spalart–Allmaras one equation turbulence model was selected for. Even though the simple RANS with the one equation turbulence model cannot simulate the highly complex U-bend flow physics precisely, the optimisation process was able to identify an optimum U-bend configuration which achieved a 63.3% pressure loss reduction, relative to the datum configuration, yielding the lowest loss U-bend in the literature. The authors also performed careful experiments to confirm their predictions and the performance of the optimum U-bend configuration identified by this work was validated.

In industrial fields of machine and aerospace, cooling systems consisting of channels are widely used to increase energy efficiency and prevent system overheat. In cooling channels, a reduced pressure drop, an enhanced heat transfer, and a short channel length are considered key design requirements for optimizing the total volume and weight of a system. In this work, we improved heat transfer efficiency by using milli-scale wavy structures inside the channel. By optimizing the inner structures through computational fluid dynamics analysis and Taguchi method, the Nusselt number increased by approximately 11.7% with a similar pressure drop compared with that of a normal channel for a Reynolds number of 1000.

The present work performs three-dimensional flow calculations based on Reynolds Averaged Navier-Stokes (RANS) and Delayed Detached Eddy Simulation (DDES) to investigate the flow field of a transonic rotor (NASA Rotor 37) at near-stall condition. It is found that the DES approach is likely to predict well the complex flow characteristics such as secondary vortex or turbulent flow phenomenon than RANS approach, which is useful to describe the flow mechanism of a transonic compressor. Especially, the DES results show improvement of predicting the flow field in the wake region and the model captures reasonably well separated regions compared to the RANS model. Besides, it is discovered that the three-dimensional vortical flows after the vortex breakdown from the rotor tip region are widely distributed and its vortex structures are clearly present. Near the rotor leading edge, a part of the tip leakage flows in DES solution spill over into next passage of the blade owing to the separation vortex flow and the backflow is clearly seen around the trailing edge of rotor tip. Furthermore, the DES solution shows strong turbulent eddies especially in the rotor hub, rotor tip section and the downstream of rotor trailing edge compared to the RANS solution.

We propose a simple method to generate hierarchical wrinkles on a single-layered substrate by using a weak photopolymerization and a thermal curing process. Rough and relatively large-scale wrinkles having 300–500 µm width sizes are fabricated by a repetitive volume dividing (RVD) method, which is known as one of the effective processes to generate wrinkle patterns in a large area. After generation of the primary wrinkles that are weakly polymerized, microscale wrinkles with width of 10–20 µm on the large-scale ridges are formed spontaneously by room-temperature-based thermal curing. We rationalize the mechanism for the generation of hierarchical abnormal shaped structures and quantify the experimental findings by a parametric study. Through this work, we show the relevance of these structures for use in diverse applications such as surface modification for wettability control and antifouling.

Erratum to “Transient performance prediction of an axial compressor considering VIGV operation speeds” [JMST 28 (10) (2014) 4099~4107] Donghyun Kim, Sangjo Kim, Changmin Son, Kuisoon Kim, Myungho Kim and Seongki Min Department of Aerospace Engineering, Pusan National University, Busan, 609-735, Korea School of Mechanical Engineering, Pusan National University, Busan, 609-735, Korea The 4th R&D Institute, Agency for Defense Development, Daejeon, Korea

In the present study, the influence of an air–oil heat exchanger (SAOHE) location and orientation on engine performance is investigated using numerical predictions with a range of geometry options that match experimental data. The airflow in the unit-fin domain of a SAOHE was modeled with rotational periodic boundary conditions. Using the standard k–ɛ turbulence model, the compressible Reynolds-averaged Navier–Stokes (RANS) equations and energy equation were solved numerically. In order to validate the numerical method, experimental measurements of the velocity profile and the distribution of the pressure and heat transfer coefficient were compared to numerical results. The pressure drop, overall heat transfer coefficient, and velocity profile downstream of the heat exchanger were taken into consideration as performance metrics. An efficient numerical procedure for the installation study of a cooler having a bypass duct was conducted, and important design variables for SAOHE were clearly identified.

In the present study, transient performance and flow characteristics of a 1.5-stage compressor with VIGV are investigated using CFD. The level of accuracy of CFD approach is evaluated by comparing the predicted performance against experimental data of a 3.5-stage compressor. The main objective of the present research is to quantitatively estimate the performance of transonic axial-flow compressor under transient operation of VIGV compared to steady-state performances on static IGV condition. The present study is performed to investigate the effect of VIGV movement on the performance characteristics when the VIGV setting angle varies with time. The VIGV is set to rotate in the clockwise and counterclockwise directions. The results were compared with steady-state predictions. Reasonable differences are observed in mass flow rate predictions (2%∼6%) between the transient and the steady-state calculations as VIGV operation speed varies. As VIGV operation speed increases, the difference in mass flow rate also increases. Therefore, a change in work input of the compressor is confirmed due to changes in mass flow rate. This is mainly caused by the changes in deviation and incidence angles depending on the operation speed and operation direction (clockwise and counterclockwise). The changes in deviation and incidence angles reflected on the entropy distribution. Even in mid-span, there is variation in angles depending on the operation speed and direction.

We numerically investigated the behavior of the water meniscus formed between a flat surface and a tip surface, which is flat or circular in shape, using the lattice Boltzmann method (LBM). The shape of the water meniscus formed between the flat bottom surface and the tip surface depends on the tip shape and the interaction between the water meniscus and the bottom or tip surface. The interaction is determined by the contact angles of the bottom and tip surfaces, resulting in different contact lengths between the water meniscus and the bottom or tip surface. The difference in these contact lengths depends on the effects of both the tip curvature and the interaction between the water meniscus and the bottom or tip surface. We classified the shapes of the water meniscus into seven different patterns as a function of the contact angles of the flat bottom and tip surfaces: concave, semi-concave, inverse semi-concave, column, convex, semi-convex, and inverse semi-convex.

A tube type heat exchanger is often the only solution when minimum pressure loss is a requirement. In addition, small diameter tubes are preferable because of an increased heat transfer area within an acceptable pressure loss limit. The present work reports on both an analytic model and experimental results with regards to the pressure drop and heat transfer characteristics of compact straight, C-curved, and U-curved tubes. The inner diameter of the tube (D) for our selected heat exchanger type was 1.26 mm with a thickness of 0.12 mm and a total length of 150.8 D. For the experiment, pressurized nitrogen gas bottles were used rather than an air compressor system in order to simplify the test facility. Hence the pressure conditions were easily set at 10, 30, and 50 bar corresponding to a range of Reynolds numbers from 10000 to 50000. To elevate the air temperature outside the tube (from 100°C to 400°C), an electric furnace was installed around the “test tube”. An analytic model to determine the pressure loss through curved tubes-referred to as the modified friction factor- is proposed. Good agreement was found between the modified friction factor and existing correlations, thus confirming the suitability of this model for determining pressure losses for different shape of tubes. The average measured Nusselt numbers were within 10- 15% of the Dittus-Boelter and Gnielinski correlations.

This study examined numerically the natural convection induced by a temperature difference between a cold outer inclined square enclosure and a hot inner circular cylinder. A two-dimensional solution for natural convection was obtained using the finite volume method with second-order accuracy and the immersed boundary method to handle efficiently the inner circular cylinder within an inclined square enclosure. The present study considered the effects of the following parameters on fluid flow and heat transfer in an enclosure: Rayleigh number from 10 3 to 10 6 , the dimensionless cylinder radii from 0.1 to 0.3 and tilted angle of the enclosure from 0° to 45°. The results showed that the distribution of isotherms, streamlines, local and surface-averaged Nusselt numbers are determined by the combined effects of convection and the distance between the cylinder and walls of the enclosure, which are a function of the Rayleigh number, dimensionless cylinder radius and tilted angle of the enclosure.

Due to its high volume goodness and light weight, the cross-corrugated primary surface heat exchanger is a promising candidate for application to an advanced intercooled-cycle gas turbine engine. The asymmetric corrugation profile considered in this paper has the possibility of achieving balanced pressure drops on both hot and cold sides of an intercooler matrix, which may enable significant weight reductions not only in counter-flow designs, but also in cross-flow designs. For the assessment of its performance, a numerical procedure incorporating a detailed three-dimensional computational fluid dynamics model and a lumped parameter flow-network model has been investigated. By developing a new parameterization method for the shape of the profile, the detailed aero-thermal performance of the asymmetric profile is summarized quantitatively. A cross-flow heat exchanger matrix sizing study for matrix weight minimization is carried out, and the results show the high dependency of the potential improvements on the given operating conditions and design requirements.

This study presents an investigation of a novel cross-corrugated primary surface for use in a heat exchanger and focuses on the design optimization of an anti-phase secondary corrugation which showed potential to improve the aero-thermal performance of a cross-corrugated heat exchanger in the accompanying paper Doo et al. [15] . In the present study, three-dimensional numerical simulations using steady incompressible Reynolds-averaged Navier–Stokes (RANS) equations with a low-Reynolds number k–e turbulence model were carried out for a range of parametric variations of the anti-phase secondary corrugation on a cross-corrugated geometry. The mechanism of enhancing the aero-thermal performance was subsequently investigated by careful analysis of the flow structures and thermal field as well as the analysis of the entropy generation in the flow passage.

Detailed heat transfer coefficient distributions have been measured on both surfaces of the impingement plate of an engine-representative impingement cooling system using the thermochromic liquid crystal (TLC) transient technique. The color images of the TLC on the impingement downstream surface provide evidence of a re-impingement flow. The reimpingement flow is found to contribute to local increases in the heat transfer on the impingement plate downstream surface. It was found that the average heat transfer coefficient on the impingement downstream surface is about 50% of the average target surface heat transfer coefficient. The results are compared with a previously reported correlation. Copyright © 2001 by ASME.

In the present study, a multi-stage transonic compressor has been analyzed to investigate secondary loss structures and flow interactions in the corner region where the hub endwall and blade suction surface meet. The Detached Eddy Simulation (DES) approach is used successfully with the Shear Stress Transport (SST) turbulence model to directly resolve the eddy structure in the separated region. The SST-DES results for a transonic three stage axial compressor are compared with a RANS analysis obtained using ANSYS CFX. The present analysis indicates that the DES is better in simulating secondary losses and vortex structures than the RANS. With the DES, a large three-dimensional separation is predicted in the stator suction surface and hub endwall compared to the RANS prediction. The flow separation affects adversely the loss characteristics such as increases in the entropy and total pressure loss. The DES analysis indicates that the secondary flow phenomenon of the stator rows is apparent in all stages. It is observed to predict two distinct vortices induced by a three dimensional flow separation in the region adjacent to the suction surface and trailing edge of the last stage stator near the hub endwall. For the front two stages, the DES also predicts strong vortices and flow separation in the same corner region while the RANS analysis fails to predict them clearly. The total pressure loss prediction is concerned, the DES analysis predicts significantly larger than the RANS analysis in the region where the hub corner separation occurs. The DES is also found to predict a periodic fluctuations in the entropy, leading to the instantaneous efficiency variations with maximum differences of about 10% compared with the RANS solutions.

The flow leaking over the tip of a high pressure turbine blade generates significant aerodynamic losses as it mixes with the mainstream flow. This study investigates the effect of blade tip geometries on turbine performance with both steady RANS and unsteady URANS analyses. Five different squealer geometries for a high pressure turbine blade have been examined: squealer on pressure side, squealer on suction side, cavity squealer, cavity squealer with pressure side cutback, and cavity squealer with suction side cutback. With the case of the cavity squealer, three different squealer wall thickness are investigated for the wall thickness (w) of 1x, 2x and 4x of the tip gap (G). The unsteady flow analyses using CFX have been conducted to investigate unsteady characteristics of the tip leakage flow and its influence on turbine performances. Through the comparison between URANS analyses, detailed vortex and wake structures are identified and studied at different fidelities. It is found that the over tip leakage flow loss is affected by the tip suction side geometry rather than that of the pressure side geometry. The unsteady results have contributed to resolve the fundamentals of vortex structures and aerodynamic loss mechanisms in a high pressure turbine stage.