bluuebunny/crossref_metadata_2025 huggingface dataset:
LES and CDNS Investigation of T106C Low Pressure Turbine
Authors: Hu Site , Zhou Chao , Xia Zhenhua , Chen Shiyi
Published: September 2016

This paper investigates the aerodynamic performance of a low pressure turbine, namely the T106C, by large eddy simulation (LES) and coarse grid direct numerical simulation (CDNS) at a Reynolds number of 100,000. The existing experimental data is used to validate the results. The effects of sub-grid scale (SGS) models, mesh densities, computational domains and boundary conditions on the result are studied. A separation zone appears downstream 60% S/S0, which is difficult to be predicted precisely. Two SGS models, i.e. Smagorinsky and WALE model are implemented. WALE model produces a better result than Smagorinsky model when the mesh is coarse. For the fine mesh, the difference due to the SGS models becomes smaller. It is surprising to find that CDNS produces very similar results to WALE model, and their results are in better agreement with the experiment than that of Smagorinsky model. The computational domain is also important for the prediction of the flow near the midspan. 3D features of the separated flow have an effect on the downstream flows, especially for the area near the reattachment. Sufficient long spanwise computational domain is important for an accurate prediction. The current study also find that the endwall secondary flow has an effect on the flow separation near the middle span, which was seldom discussed before. By considering the effects of endwall secondary flows, better prediction of the flow separation near the blade midspan can be achieved. The effect of the endwall secondary flow on the blade suction side separation is also explained with the analytical method based on the Biot-Savart Law.
An Experimental Investigation of Full-Coverage Film Cooling Effectiveness and Heat Transfer Coefficient of a Turbine Guide Vane in a Linear Transonic Cascade
Authors: Fu Zhong-yi , Zhu Hui-ren , Liu Cun-liang , Liu Cong , Li Zheng
Published: September 2016

This paper experimentally investigates the film cooling performance of an enlarged turbine guide vane with full-coverage cylindrical hole film cooling in short duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and isentropic exit Mach number independently. The effects of mass flow rate ratio (MFR=4.83%∼8.83%), inlet Reynolds number (Rein= 1.7×105∼5.7×105), and isentropic exit Mach number (Mais=0.81∼1.01) are investigated. There are five rows of cylindrical film cooling holes on the pressure side and four such rows on the suction side respectively. Another four rows of cylindrical holes are provided on the leading edge to obtain a showerhead film cooling. The surface heat transfer coefficient and adiabatic film cooling effectiveness are derived from the surface temperatures measured by the thermocouples mounted in the middle span of the vane surface based on transient heat transfer measurement method. Mass flow rate ratio is shown to have a significant effect on film cooling effectiveness. The increase of mass flow rate ratio increases film cooling effectiveness on pressure side, while increasing this factor has opposite effect on film cooling effectiveness on the suction side. At the same mass flow rate ratio, increasing the Reynolds number can enhance the film cooling performance, the expectation is that at low mass flow rate ratio condition increasing the Reynolds number decreases film cooling effectiveness on the pressure side. The heat transfer coefficient increases with the mass flow rate ratio increasing on both pressure and suction side. At middle and high inlet Reynolds number condition, in the region of 0.4<s<0.6 on suction side, the coolant weakens heat transfer adversely.
Assessment of Grid Resolution Requirements for Accurate Simulation of Disparate Scales of Turbulent Flow in Low-Pressure Turbines
Authors: Pichler R. , Sandberg R. D. , Michelassi V.
Published: September 2016

Direct numerical simulations (DNS) on different grids for two Reynolds numbers of the T106A turbine cascade have been conducted to understand the influence of resolution on the solution. The work was motivated by different resolutions used in previously published studies that could not only be explained by different numerical methods. Most of these published studies used mean blade pressure distribution and wake loss profiles for validation and grid convergence purposes. DNS for canonical cases in general use more stringent criteria based on turbulence statistics considering that those require higher resolution.

Therefore, in this study the influence of grid resolution on disparate scales of DNS is discussed. It is discussed that mean blade pressure distribution and mean wake loss already show good agreement with experiments for large eddy simulations (LES) provided a sufficient part of the turbulence kinetic energy is resolved. However, to avoid modelling uncertainties of turbulence quantities, DNS is required as LES only resolves part of the turbulence kinetic energy. It is found that in order to achieve grid convergence for turbulence dissipation a six times larger resolution and more than 10 times the computational cost is required than if grid convergence was required for turbulence kinetic energy.
An Introduction of Aerodynamic Shape Optimization Platform for Compressor Blade
Authors: Yanhui Duan , Wenhua Wu , Zhaolin Fan , Ti Chen
Published: September 2016

In this paper, an aerodynamic shape optimization platform for compressor blade is introduced. The platform divided into modules on flow field calculation, optimization method, parameterization and grid deformation. Flow field calculation of compressor blade is based on computational fluid dynamics (CFD), which is used for multi-block structure grid. Particle swarm optimization (PSO) is built as optimization method module, in which cost functions are calculated parallel. In parametric module, 3D blade is decomposed in a series of characteristic sections and the section is parameterized by Hicks-Henne function. Algebraic interpolation method is used for grid deformation, which is a high efficiency and robust method. Two cases of rotor 37 are presented. The result of the first case shows that, the CFD code of the optimal platform is reliable and robust. For the second case, the optimal platform is verified by designing rotor 37. The result shows that, the optimal platform is effective for design of compressor blade.
Unsteady and Multi-Blade Passage Numerical Simulation on an Unshrouded Centrifugal Impeller
Authors: Zhang Tongtong , Wang Tong , Wang Hui , Gu Chuangang
Published: September 2016

Unsteady numerical simulation was performed with both single blade passage model and 4-blade passage model with ANSYS CFX software. Grid independence and turbulence model were verified by comparison with the experimental results. By transient numerical simulation, the time dependent parameters in tip area were analyzed. The typical vortex region was found near the blade suction side by relative vorticity strength. The pressure distributions at 90% span of the blades were compared. The mixing of tip leakage flow and the main flow was studied. 4-blade passage model simulation can capture circumferential difference in the centrifugal impeller. Periodic flow structures can be observed. Frequency analysis was performed on the pressure fluctuation in the blade tip. The effectiveness of 4-blade passage model in capturing the unsteady flow structures in the centrifugal impeller was demonstrated. The conclusion can provide guidance for stability analysis on the centrifugal impeller.
Flow Pattern Evolution and Energy Decomposition of Flows at Different Operating Conditions in a Hydrodynamic Torque Converter
Authors: Wei Wei , Huang Mingxing , Li Yu , Yan Qingdong
Published: September 2016

Power loss and flow blockage in turbomachinery such as hydrodynamic torque converter are usually caused by jet flow, second flow and flow separation. In this paper, the velocity vector and the pressure distribution of the internal flow field in hydrodynamic torque converter were reduced by the method of the Proper Orthogonal Decomposition (POD) to find the main flow structures and the energy decomposition in the passages of pump, turbine and stator. In order to find their evolutionary processes and energy decompositions, oil flow visualizations were conducted at different speed ratios from 0 to 0.8, including stall condition and design operating condition.

The results showed that the first few modes containing the majority of energy could provide enough accuracy to predict flow behavior and flow structure in flow passages. Especially when the energy percentage of the first mode was majority, its vortex structures could be recognized easily. But the flow patterns of other modes were different from each other and they made the flow more turbulent and complex, which increases the energy loss in the process of power transmission. Besides that, the change of pressure gradient had a direct influence to velocity vector. The results also indicated that the observed fluid pattern of vortex structure became extensive while the influence of secondary flow decreased in the flow passage of pump with the increase of speed ratio. But the situation is just reversed in turbine, that is, the vortex disappeared gradually and the irregular turbulent flow appeared as the increase of speed ratio. In stator, the vortex structure emerged gradually when the speed ratio increased. So the method of snapshots is a very useful way to analyze the complex flow flied in depth and to predict the trend of development.
3D Unsteady Simulation of a Low Speed Low-Reaction Aspirated Compressor
Authors: Zhang Longxin , Du Xin , Liu Xun , Wang Songtao
Published: September 2016

Aspirated compressor is a promising design concept to promote the working capacity of compression system and thus may result in an intensive interactions with adjacent rows, however, few extensive unsteady flow analysis have been conducted on this type of highly-loaded compressor. This work presents a three-dimensional numerical simulation of a low speed 1.5 stage low-reaction aspirated compressor (LRAC). The boundary layer suction is only implemented in the outlet vane to control the corner stall and boundary layer separation while no active flow control method is applied in the rotor. The total aspiration flow rate is around 3%. Both aspiration slot and plenum were integrated into the computational domain. Two operating points were selected with the aim to investigate the unsteady effects on the performance of the LRAC and to provide a preliminary unsteady description for this type of aspirated compressor. It is found that compared with the differences in 1D total aerodynamic parameters, evident departures are found in the radial distributions of stage outlet flow parameters between the steady and time-averaged results. For the unsteady case, the radial distributions of pitchwise-averaged parameters become more uniform due to the redistributed aspiration flow rate and the convection behavior of the rotor wake. For the aspiration scheme in the blade, although one-side-aspiration manner is applied, the aspiration flow rate presents a C-type distribution in the radial direction, and this tendency becomes more prominent for the forward slot and time accurate results. Besides, the fluctuation of aspiration flow rate is mainly focused on the upper span due to the intensive rotor outlet secondary flow here. Moreover, the potential effect of aspirated stator on rotor is also examined preliminarily. It is found that for the LRAC investigated in this paper, the rotor suction side is apt to be influenced by the downstream aspirated stator. Finally, some suggestions on the design of the aspirated compressor are provided.
Aerodynamic Design Optimization of the Last Stage of a Multi-Stage Compressor by Using an Adjoint Method
Authors: Yang Jing , Luo Jiaqi , Xiong Juntao , Liu Feng
Published: September 2016

The paper presents the aerodynamic optimization of the last row of a 4.5-stage compressor by means of a gradient-based optimization method that utilizes the continuous adjoint approach. An adjoint mixing-plane formulation is used to allow for the computation of adjoint solutions for multistage turbomachinery problems. Firstly, a conventional one-dimensional design method and empirical correlations are used to produce a base design of a 4.5-stage low-speed, low compression ratio compressor with an inlet guide vane to adjust to different operation conditions. Then the last stage is redesigned by the adjoint method to reduce the flow losses at the operation condition near stall through modifying the aerodynamic shape and stagger angle of the stator blade. The cost function is defined as a weighted sum of the entropy production and a constraint on the mass flow rate. Finally, a multi-point design optimization approach by using the adjoint method is employed to improve the performance of the last stage at two different operation conditions. The results demonstrate that the stator shape modifications improve the aerodynamic performance of the stage and illustrate the functionality of the adjoint-based multi-stage optimization system.
Automated Multi-Code URANS Simulation of Compressor-Combustor Components
Authors: Kannan K. V. , Page G. J.
Published: September 2016

Currently in an aircraft gas turbine engine, the turbomachinery and combustor components are designed in relative isolation and the effect of the upstream and downstream components on each other’s flow are not fully captured in the design process. The objective of this work is to carry out a multi-code integrated unsteady simulation of Compressor-Combustor components with each zone simulated using its own specialised CFD flow solver. The multi-code URANS technique is simple, based on files and involves the generation of new 2D boundary conditions for the required flow field at each time step. A driver based on a Python script automates the entire process. This paper shows the method first validated in a simple vortex shedding 2D case and then extended to a cold flow URANS simulation matching an isothermal compressor/combustor rig experiment. An external coupler code is invoked that produces unsteady, spatially varying, inlet conditions for the downstream components. The simulation results are encouraging as the mass, momentum and energy losses across the interface are less than 1%. The multi-code unsteady simulation produces wake profiles closer to the experiment than the coupled steady RANS simulation. The present study shows a reasonable agreement with the experimental PIV and hot-wire data thus demonstrating the potential of the multi-code integrated simulation technique.
Design and Optimization of Tandem Cascade Based on Parallel Differential Evolution Algorithm
Authors: Cheng Hao , Liu Bo , Yang Xiaodong , Li Jun
Published: September 2016

A method of tandem airfoil geometry generation and an algorithm of master-slave parallel differential evolution are first developed for the later optimization. An initial tandem cascade is roughly designed and significantly outperforms the original conventional cascade. Based on the parallel differential evolution algorithm and a Navier-Stokes solver, five configuration variables of the initial tandem cascade are then numerically optimized at an inlet Mach number of 0.7 and an approximately minimum-loss incidence of 1.9°. The result shows that the total pressure loss coefficient of the optimum design decreases by 8.67%. The history data of the optimization is statistically analyzed, which reveals the influence levels of the five configuration variables on tandem performance. The performances of the initial and optimum designs at a range of incidence angles are then numerically calculated, showing that the optimum design outperforms the initial design at small or negative incidence angles and performs more poorly at high incidence angles. It is proposed and verified that the different front-rear distributions of camber and chord leads to this phenomenon. Finally, a new-defined variable is proposed to measure the distribution above.
Flow Analysis and Uncertainty Quantification of a 2D Compressor Cascade With Dirty Blades
Authors: Qin Ruihong , Ju Yaping , Wang Yazhou , Zhang Chuhua
Published: September 2016

Compressor cascades with service inevitably suffer from blade fouling or erosion, which modifies blade geometries and brings sufficient damage to compressor performance. However, very few studies were cast upon the impact of degraded blade surface morphologies on the compressor aerodynamic performance. In this study, a novel geometry parameterization method is proposed to describe the non-uniform, multi-scale and stochastic geometric features of blade surface morphologies in fouling and erosion conditions. Through computational fluid dynamics (CFD) simulations, the influences of coverage, dirty size and density of dirty units on the cascade performance and flow field are numerically investigated. Finally, uncertainty analyses are conducted to quantify the cascade performance variations due to geometric uncertainties of dirty surface morphologies. To alleviate the computational load, a relatively new dimension reduction technique, i.e., the active subspace method is used to reduce the original eight-dimensional uncertainty quantification problem to a one-dimensional one. The CFD results show that at the high flow rate, the cascade performance is gradually worsened as the dirty surface morphology extends from the blade leading edge to the trailing edge on the suction surface. However, at the low flow rate, the cascade performance degradations tend to be alleviated once the dirty surface morphology enters the region of the separation bubble on the suction surface. At both operating points, the cascade total pressure loss almost increases monotonically with either larger or densified dirty units distributed on the blade suction surface. The probabilistic distributions from uncertainty analyses show that the cascade performance is more sensitive to the dirty surface morphology at low flow rates than at high flow rates. The present study can help understand the loss mechanisms within the 2D cascade in fouling and erosion conditions, which is beneficial for the enhancement of 3D compressor performance.
Investigation on the Turbine Blade Tip Clearance Monitoring Based on Eddy Current Pulse-Trigger Method
Authors: Wang Weimin , Shao Huajin , Chen Lifang , Song Huibin
Published: September 2016

The efficiency and reliability of turbomachinery will be improved by blade tip clearance (BTC) and blade tip timing (BTT) monitoring. Several types of sensors such as eddy-current, capacitance and optical probes are used to realize this objective. Eddy current sensor (ECS) is an ideal choice with its advantage of durablity and that it is unaffected by gas stream properties such as contamination, water vapor, and moisture. However, the bandwidth of ECS is usually less than 100 kHz, which will limit the resolution of the monitoring result. In this paper, a pulse-trigger technology based BTC method was presented. This method optimizes the static radial and circumferential calibration technology to obtain the sensitivity of the ECS in the different relative locations against the tip of blade. The information from the clearance sensor will be fused with that from the once per revolution (OPR) or key phase sensor. The method is more generally applicable in the condition where the ECS is insufficient sampling caused by the limit of narrow bandwidth, especially under the high blade tip velocity condition. A small scale and larger scale BTC measurement rig are established to validate the feasibility of this method. The small one is easy to calibrate with high accuracy and can be used to illustrate the performance of the method, while the larger scale test rig is close to real industry turbine blade. In this apparatus, the axial displacement and radial displacement of rotor vibration as well as the clearance can be monitored together so that further investigation can be conducted. Experimental research was carried out on both test rig at different rotating speed. The results show that the method presented in this paper can improve the accuracy of tip clearance monitored by ECS very well. Furthermore, this work is a proof-of-concept demonstration using a laboratory setup providing the basis for BTC active control and blade health monitoring (BHM) based on ECS.
Numerical Evaluation of the Condition of a Jet Engine Through Exhaust Jet Analysis
Authors: Adamczuk Rafael R. , Seume Joerg R.
Published: September 2016

The present work numerically investigates the influences that four defects in the hot gas path of a civil aircraft engine have on the density distribution of the exhaust jet. The defects detectability with tomographic measurements using the Background-oriented schlieren method (BOS) is evaluated in order to draw conclusions on the condition of the engine before disassembly. The modeled defects are: the variation of the radial gap between the rotor blade tip and the casing in the second stage of a high-pressure turbine (HPT), the burning of the trailing edge of the stator vanes and rotor blades of the second stage of the HPT, the variation of the cooling air mass flow in the first stage of the HPT, and the variation of the temperature at the outlet of the combustion chamber. Synthetic measurements show that a characteristic signature at the outer border of the core mass flow in the exhaust jet can be measured with BOS when the radial gap is varied. However, the burning of the trailing edges cannot be detected. The variation of the cooling air massflow affects the entire density distribution and can be measured with BOS. Defects of the burner nozzles affect a larger local area of the exhaust jet and thus can also clearly be identified with BOS. The results show that different defects in the hot gas path result in characteristic signatures in the density distribution of the exhaust jet and that they can be measured with BOS. It is thus possible to detect defects in the hot gas path with BOS before the disassembly of the engine.
Effect of Quarl on N2- and CO2-Diluted Methane Oxy-Flames Stabilized by an Axial-Plus-Tangential Swirler
Authors: Jourdaine Paul , Mirat Clément , Beaunier Jérôme , Caudal Jean , Joumani Youssef , Schuller Thierry
Published: September 2016

Stabilization of nearly stoichiometric CO2- and N2-diluted premixed methane/oxygen swirling flames is investigated in an atmospheric test rig equipped with an axial-plus-tangential swirler exhausting in a cylindrical injection tube eventually ended by a diverging quarl. The investigated flames are stabilized aerodynamically away from the solid elements of the combustor without the help of any central bluff body in the injector. The flowrates through the axial and tangential slits of the swirler can be adjusted separately. Effects of swirl and quarl angle on the flowfield and flame shape are analyzed. Laser tomography on small oil particles reveals that fuel, oxidizer and diluents injected in separated channels are well mixed at the injector outlet. Velocimetry measurements and flame images show that the axial-plus-tangential swirler allows a flexible control of the flame leading edge position with respect to the injector outlet. For a fixed injector geometry with a given quarl angle and swirl number, it is found that N2- and CO2-diluted flames feature the same topology if the injected combustible mixtures feature the same adiabatic flame temperature, while they may feature different bulk injection velocities and laminar burning velocities. The operability range of well stabilized CO2-diluted flames is however reduced.
Impingement/Effusion Cooling Wall Heat Transfer: Conjugate Heat Transfer Computational Fluid Dynamic Predictions
Authors: El-Jummah Abubakar M. , Andrews Gordon E. , Staggs John E. J.
Published: September 2016

Impingement/effusion internal wall heat transfer was investigated using conjugate heat transfer (CHT) computational fluid dynamics (CFD). ANSYS Fluent was used with the k-ε turbulence model. Impingement/effusion configurations were modelled with the impingement hole on the centre of the effusion hole square array. Impingement/effusion geometries with equal number of holes were modelled with hole density n (m2) of 4306/4306, 9687/9687 and 26910/26910 m−2 for an impingement hole pitch to diameter ratio X/D of ∼ 11 at an effusion hole X/D of 4.7 for all configurations. The experimental test wall size was 152mm square this was square hole arrays of 102, 152 and 252. The main reason for investigating the number of holes was that effusion cooling has been shown to improve the greater the number of holes used for the same wall porosity or X/D. The coolant mass flux G was varied from 0.1–0.94 kg/sm2bar for all n. A constant impingement gap Z of 8 mm was used with both walls 6.35 mm thick. The range of gap to diameter ratio Z/D was from 2.4–12.5, with the smallest n having the lowest Z/D. Locally surface average heat transfer coefficient (HTC) h values showed reasonable agreement with previously published experimental results. Comparison of the h predictions on the effusion (target) approach wall with those for the effusion and impingement walls alone, were made and this showed that impingement/effusion wall heat transfer was less than the sum of the two components. The reason was shown to be due to the reduced internal gap flow recirculation with reduced heat transfer to the impingement wall for impingement/effusion cooling.
Solid Element Rotordynamic Modeling of a Rotor on a Flexible Support Structure Utilizing MIMO Support Transfer Functions
Authors: Hu Lingnan , Palazzolo Alan
Published: September 2016

The accurate modeling of a rotor system is essential for effective design and troubleshooting in rotating machinery. The beam-type finite element (FE) may be inadequate for modeling a rotor or support structure with complex shapes. In addition, the isolated support impedance methods may be inaccurate for modeling the support structure that has modes that are highly coupled between bearings and directions at the bearing locations. The solid FE method is a good replacement of the beam FE and support impedance approaches. However, a drawback for this method is the significant amount of computation time required to obtain accurate solutions due to the large number of nodes in the solid FE analysis.

The authors present an improved approach to analyze the coupled rotor-support dynamics, by modeling the rotor with solid elements and utilizing transfer functions (TFs) to represent the flexible support. A state-space model is then employed to perform general rotordynamic analyses. The solid FE rotor model includes the gyroscopic effects and the asymmetric and cross-coupled stiffness coefficients of the bearing. A series of rational TFs are used to simulate dynamic characteristics of the support structure, including the cross-coupling between degrees of freedom (DOFs). These TFs are derived by curve-fitting the frequency response functions (FRFs) of the solid FE support model at the bearing locations. The impact of the polynomial degree of the TF on the unbalance response analysis is discussed, and a general rule is proposed to select an adequate polynomial degree.

To validate the proposed modeling approach, a comprehensive comparison among the complete solid FE rotor-support model and the solid FE rotor model with the TFs representing the flexible support (the reduced state-space model) are presented. Comparisons are made between natural frequencies, critical speeds, unbalance response, logarithmic decrement (log dec), and computation time. The results of these comparisons show that the reduced state-space rotor-support model provides a dynamically accurate approximation of the solid FE rotor-support model in terms of general rotordynamic analyses. Moreover, the computation time for the proposed modeling approach is reduced to 2.5 minutes, compared to 14 minutes for the complete solid FE modeling. The reduction of the computation time may vary with different number of DOFs of the rotor model and the support structure model.

In addition, the modes up to 100,000 cpm are compared among the beam rotor with the solid FE support model, the solid FE rotor with the super-element support model, and the reduced state-space model. The results show that the reduced state-space model is more accurate in predicting high-frequency modes than the beam rotor-support and super-element support models. Further, the proposed approach with the state-space model is useful for applications in vibration control and active magnetic bearing (AMB) systems.
Assessment of Body Force Methodologies for the Analysis of Intake-Fan Aerodynamic Interactions
Authors: Thollet William , Dufour Guillaume , Carbonneau Xavier , Blanc Florian
Published: September 2016

With the advent of high bypass ratio turbofan engines of increasing size, new nacelle designs with shorter air intakes have to be considered, creating new aerodynamic interactions with the fan. This paper focuses on a numerical strategy known as Body Force Modeling, in which engine components are modeled and taken into account using source terms in the RANS equation. This approach allows to simulate these interactions with an accuracy comparable to full 360° unsteady simulations, but at a fraction of the cost. Different formulations for the body forces are proposed and applied to different nacelle test cases of varying intake length. Using full annulus unsteady computations as validation data, it is shown that the body force approach allows to capture inlet design effects on in-plane forces and on nacelle flow separation.
Flow Characteristics of an Optimized Axial Compressor Rotor Using Smooth Design Parameters
Authors: Mahmood Syed Moez Hussain , Turner Mark G. , Siddappaji Kiran
Published: September 2016

Blade designs have evolved from NACA series and free vortex assumptions to detailed meanline and forced vortex definitions. A design process is presented with numerous parametric options to explore a large design space. Smoothness in turbomachinery blade shapes is critical to an effective design. A cubic B-spline is used to control spanwise variations in the curvature definition of airfoil camber, thickness distribution, leading edge definition, inlet angles and outlet angles as parameters with a small number of control points. Varying parameters of individual blade sections requires more control variables that increases the parameter space and adds kinks in the 3D blade shape. Benefits of this smooth spanwise capability are demonstrated by linking the blade design tool with an aerodynamic optimization system. A single subsonic rotor (rotor 6 of a 10 stage axial compressor derived from the GE EEE design) has been considered as the baseline for the optimization process. Optimization is performed by varying curvature of the airfoil camberline as well as inlet and outlet angles in the spanwise direction. A single objective optimization was performed to optimize isentropic efficiency. An improvement in efficiency of 0.83% from 91.87% to 92.63% was obtained. The optimized blade geometry has a smooth transition from a traditional airfoil shape at the hub section to an S-shaped airfoil at the mid and tip sections. This unique blade shape was obtained because the airfoil camber curvature definition was allowed to vary smoothly spanwise. An S-shaped blade near the mid and tip section promotes flow to move radially downwards which allows for a reduction in entropy generation due to tip leakage flows. Entropy is used to quantify losses and improvement in efficiency.
Area-Schedule Based Design of High Pressure Recovery Radial Diffusion Systems
Authors: Gao Ruhou , Spakovszky Zoltán , Rusch Daniel , Hunziker René
Published: September 2016

High-pressure ratio centrifugal compressors require advanced diffusion systems to achieve enhanced efficiencies set by future turbocharger applications. To address the shortcomings of the commonly used channel diffuser and airfoil cascade design perspectives, a streamtube based area-schedule is adopted paying special attention to the diffuser entry region. It is shown that the diffusion in the semi-vaneless-space, controlled chiefly by inlet flow angle and the vane suction side geometry, plays a key role in improving diffuser performance. Removing excess thickness from the suction side eliminates flow overspeed, increases effective diffusion length, and leads to higher pressure recovery at reduced stagnation pressure loss. The pressure side thickness distribution controls the channel area schedule. Thin leading edges ensure smooth flow area transition into the channel and reduce the vane upstream influence, mitigating high-cycle fatigue related mechanical issues.
Unsteady Heat Release of Stratified Swirling Flames in a Partially Premixed Pre-Vaporized Combustor
Authors: Wang Jing , Hui Xin , Qin Hao , Zhang Chi , Lin Yuzhen
Published: September 2016

Unsteady heat release of stratified swirling flames in a partially premixed pre-vaporized combustor at elevated pressure and temperature is experimentally investigated. Stratified flames produced by a pilot swirler and a main one which are fueled by different flow rates of aviation kerosene. Three fuel staging ratios (SR) of 25%, 30%, and 35% are tested to study its effect on the heat release of stratified flame. CH* chemiluminescence emissions are recorded by a high speed ICCD camera with a 430±10 nm band-pass filter. The Proper Orthogonal Decomposition (POD) analysis is employed to identify the most prominent energetic modes of flame fluctuations. Test results show that the low frequency fluctuation of the flame global heat release is weakened as the SR decreases while the high frequency fluctuation is enhanced. According to the analysis of average heat release of the flame, SR affects the structure of flame by changing the flame angle. The stratified flame dynamics mainly performs the bulk motion, asymmetric motion, and flame shedding. It is found that the flame bulk motion modes contain higher energy than those of other modes. While the other two lower energy modes would transit from asymmetric motion to flame shedding mode as the SR decreases. From the Fast Fourier Transform (FFT) analysis, the low frequency fluctuation of global heat release is dominated by modes of higher energy, while the high frequency fluctuation is related to all the modes.
Numerical Investigation of the Flow Distortion Impact on a Refrigeration Centrifugal Compressor
Authors: Shen Feng , Yu Lei , Cousins William T. , Sishtla Vishnu , Sharma Om P.
Published: September 2016

The inlet pipe to a centrifugal compressor in a water-cooled chiller can cause complexities in designing a compact unit. Flow distortion caused by the inlet pipe shape and non-uniformity of the flow from the evaporator can affect the compressor performance. Inlet guide vanes (IGVs) used for capacity control may help reduce the distortion by forcing a redistribution of the flow. However, the guide vanes themselves result in increased losses and can introduce flow distortion due to the clearance between the movable vane and stationary housing. In this effort, the influence of the flow distortion on the impeller performance and the losses caused by IGVs were investigated using a commercial code dealing with real gas. The CFD model was first verified by comparison to test data. Further investigation showed that the hub-to-shroud flow distortion, especially shroud side distortion, has a major effect on stage performance. IGVs are necessary for straightening the incoming flow to benefit stall margin but generate high losses themselves. The tip clearance of the IGV decreases the deswirl capability and introduces additional local separation loss. Distortion intensity was calculated according to ARP-1420 methods to justify the usage of the mixing plane in the CFD modeling. The extent of distortion indicated that the steady simulation served as a preferable choice for balancing computational cost & prediction accuracy.
Modeling for Tip Clearance Effects on Stall-Onset Condition in Transonic Axial Compressors
Authors: Ye Shubo , Zhao Qingjun , Xiang Xiaorong , Xi Guang , Xu Jianzhong
Published: September 2016

Rotating stall is a primary limit to compressor performance, and the reasonable estimation of stall-onset point is very useful in compressor design. Extensive investigations have been conducted in the past few decades to develop analytical models and numerical methods for stall-onset prediction, and much progress has been made in the understanding of flow mechanism for rotating stall. In spite of the robust prediction ability of stall-onset condition, the unsteady 3-D computations are still time-consuming for industrial applications. Analytical models are able to provide a fast estimation of compressor stall onset. However, empirical correlations are usually needed in the analytical models, which leads to a decrement in the accuracy and application scope of the models. Especially for high speed compressors, tip clearance effects hasn’t been evaluated reasonably in the previous analytical models, which actually plays a very important role in determining the stall-onset point. Therefore, new analytical models accounting for tip clearance effects will be promising in estimating the stall-onset more precisely. It’s the requirement for a new analytical model that motivates the present work.

In the present work, the unsteady flow simulation of a transonic compressor rotor at near stall condition was performed to clarify the relations between tip clearance flow oscillations and compressor stall-onset in transonic axial compressor rotors. The interaction between tip clearance and incoming flow is simplified to a 2-D analogy of free-stream and counter-flow wall jet interaction. Momentum balance analysis is applied to identify the position of tip clearance/incoming flow interface, together with a prediction method of tip leakage vortex core trajectory. The effects of the in-passage shock on tip clearance/incoming flow interaction is taken into account by applying an upstream deflection of the interface, and this deflection is also observed in the computational flow field at near stall conditions. As a combination of the above-mentioned aspects, a model is proposed to define the critical point for tip clearance flow spillage from blade leading edge, which corresponds to the stall-onset point on compressor performance curves. Validations against numerical results prove that the model is capable of including tip clearance effects on stall-onset point. Parametric study of the model shows that blade tip offloads with increasing tip clearance, reducing the inverse momentum flux of tip clearance flow. As a result, the stalling flow coefficient appears to be less sensitive to tip clearance variation, which accords with the published experimental results.
Pressure Field Measurement on Compressor Cascade Using Pressure-Sensitive Paint
Authors: Gao Limin , Hu Xiaoquan , Tian Linchuan , Li Ruiyu
Published: September 2016

In the current work, pressure distribution on the surface of blade and endwall is investigated experimentally in a compressor cascade for capturing some flow features by analyzing pressure profile aerodynamics. Pressure is measured using pressure sensitive paint (PSP) technique. Before the tests, some system errors including temperature-dependence of paint, in-situ calibration, and Self-illumination among neighboring structure are studied in details. In addition, the spatial layout scheme of LEDs and CCDs for each test is designed by self-developing optical theoretical method that based on the LTS (light Tracing Simulation). In the post-processing, an in-situ calibration is conducted to convert the intensity of luminescence to pressure, and each test uses independent calibration equation. Finally, discrete pressure is also measured by pressure taps to compare with PSP. All of the tests are performed at the AOA (Attack of Angle) of −2.5°∼10°, and Mach number of 0.4∼0.8. The results show that the pressure distribution at low Mach number presents a good PSP performance in terms of signal-to-noise-ratio, and apparent pressure step gradient is captured at high Mach number. Meanwhile, the pressure distribution at high incidence reveals the unsteady of separated flow. By combining the end-wall plot with profile plot, the influence of corner-flow on main flow in the passage is obtained.
Calculation of Film-Cooling Effectiveness and Aerodynamic Losses Using DES/SAS and RANS Methods and Compared With Experimental Results
Authors: Lin Xiao-Chun , Liu Jian-Jun , An Bai-Tao
Published: September 2016

When used to simulate the film-cooling internal flow, RANS models generally over predicted the mixing process of the coolant and the mainstream in streamwise, and under predicted the distribution of the coolant in spanwise. Compared with an anisotropic amendment of the RANS model, the transient LES model was shown to be universally effective to obtain more accurate prediction on mixing process. Considering the giant computing resources consumption of the LES model, this paper employs DES (Detached Eddy Simulation) SAS (Scale-Adaptive Simulation) and RANS models to study the film cooling mixing process between the coolant and the mainstream. Vortex shedding phenomenon and the detailed structure were studied, and the mechanism of shedding was exposed. Film cooling effectiveness and aerodynamic losses of simulation results were compared with those of experimental results. Result discrepancies between numerical models and experiment were discussed. It showed that DES and SAS model was more accurate than RANS on the prediction of mixing process in the spanwise and in the height directions.
Efficiency Improvements in an Industrial Steam Turbine Stage: Part 1
Authors: Deshpande Srikanth , Thern Marcus , Genrup Magnus
Published: September 2016

The present work approaches the idea of increasing the efficiency of an industrial steam turbine stage. For this endeavor, an industrial steam turbine stage comprising of prismatic stator and rotor is considered. With the velocity triangles as input, airfoil design is carried out. Firstly, the rotor is redesigned to take care of any incidence issues in the baseline case. In rotor blades, the peak Mach number is reduced in blade to blade flow passage and hence, efficiency of stage is increased. Rotor is made front loaded. After finalizing the rotor, the stator is redesigned. Stator is made more aft-loaded when compared to the baseline case. By making the stator aft-loaded, the efficiency increased by reducing profile losses. This design modification also showed advantage in secondary losses. The total pressure loss in the stator was reduced by a delta of 0.15. When creating an airfoil for stator or rotor, MISES was used in order to evaluate profile losses. The design verification for the stage was numerically done using commercial CFD software ANSYS CFX. Steady state RANS simulations were carried out. The stator and the rotor still being prismatic, only by virtue of airfoil design, the total to total stage efficiency improvement of 0.33% was predicted.
Assessment of Scale Resolved CFD Methods for the Investigation of Lean Burn Spray Flames
Authors: Andreini A. , Bertini D. , Mazzei L. , Puggelli S.
Published: September 2016

Incoming standards on NOx emissions are motivating many aero-engines manufacturers to adopt the lean burn combustion concept. However, several technological issues have to be faced in this transition, among which limited availability of air for cooling purpose and thermoacoustics phenomena that should be managed to safely implement this burning mode. In this scenario, standard numerical design tools are not often capable of characterizing such devices. Thus, considering also the difficulties of experimental investigations in a highly pressurized and reactive environment, unsteady scale resolved CFD methods are required to correctly understand the combustor performances. In the last years Large Eddy (LES) and hybrid RANS-LES models such as Scale Adaptive Simulations (SAS) have undergone considerable developments. Such approaches have been already applied for gaseous flames, leading to a strong enhancement in phenomena prediction with respect to standard steady-state simulations. However, huge research efforts are still required when spray flames are considered, since all the numerical models chosen to describe spray dynamics and the related reactive processes can have a strong impact on the accuracy of the whole simulation. In this work a set of scale resolved simulations have been carried out on the DLR Generic Single Sector Combustor spray flame for which measurements both in non-reactive and reactive test conditions are available. Exploiting a two-phase Eulerian-Lagrangian approach combined with a Flamelet Generated Manifold (FGM) combustion model, LES simulations have been performed in order to assess the potential improvements with respect to steady state solutions. Additional comparisons have also been accomplished with SAS calculations based on Eddy Dissipation combustion model (EDM). The comparison with experimental results shows that the chosen unsteady strategies lead to a more physical description of reactive processes with respect to RANS simulations. FGM model showed some limitations in reproducing the partially-premixed nature of the flame, whereas SAS-EDM proved to be a robust modelling strategy within an industrial perspective. A new set of spray boundary conditions for liquid injection is also proposed whose realiability is proved through a detailed comparison against experimental data.
Effects of Rotation on Heat Transfer due to Jet Impingement on Cylindrical Dimpled Target Surface
Authors: Singh Prashant , Ekkad Srinath
Published: September 2016

Jet impingement has been extensively used in gas turbine airfoil internal cooling, especially leading edge and mid-chord region, which are subjected to high heat transfer loads. Earlier studies have shown that impingement onto dimpled surface results in higher heat transfer augmentation, particularly in maximum crossflow setting. The present study investigates the effect of Coriolis force and buoyancy in rotating channel featuring round shape jet plate and dimpled target plate, where the spent air is allowed to exit in only one direction, thus imposing maximum crossflow. In order to ensure that the buoyancy acts in a direction similar to actual gas turbine blades, colder air is passed into the feed chamber (also in rotation). Detailed heat transfer measurements are presented using transient liquid crystal thermography, where the target surface is modeled as one-dimensional semi-infinite solid. The jet plate features three (and four) rows of circular holes, where the normalized spanwise and streamwise pitch (x/d, y/d) of jets is kept at 3. The normalized jet-to-target surface distance (z/d) is 4 and the nozzle aspect ratio is 2. The target plate features four rows of dimples, where the dimple print diameter to jet hole diameter ratio is 0.61. The cylindrical dimple arrangement on the target plate has been kept the same for all the configurations tested, where the dimple pitch to dimple print diameter ratio is kept at 1.832. The baseline case for the normalization of Nusselt number is the smooth target surface. Three different configurations of jet plate has been studied. The flow and rotation conditions have been kept the same for all the configurations, where the average Reynolds number (based on jet hole hydraulic diameter: Rej) has been maintained at 2500 and the rotational speed has been kept at 400 RPM (corresponds to Roj of 0.00274). Flow experiments have been carried out to determine the variation of discharge coefficient of each hole with the plenum absolute pressure to ambient pressure ratio, and these Cd measurements were used in the determination of the flow distribution in the jet hole plate. Under non-rotating conditions, dimpled target surface have heat transfer enhancements greater than unity. Also, for non-rotating cases, crossflow effects have been seem to be maximum and for long exit lengths, dimpled target surfaces might have lower heat transfer as compared to smooth target. It has been found that rotation has negative effects on heat transfer augmentation on dimpled target surface for all three configurations studied.