bluuebunny/crossref_metadata_2025 huggingface dataset:
Development of a Novel Axial Compressor Generation for Industrial Applications: Part 1 — Compressor Design and Performance
Authors: Ziegler Kai , Eisenberg Bernhard , Hörmeyer Klaus , Emmrich Roland , Ressing Henning
Published: September 2014

An industrial axial compressor has to meet a wide range of operation requirements and therefore must run within the whole compressor map without restrictions at an overall high level of efficiency. Additionally a robust design is required allowing a continuous operation of up to five years under industrial boundary conditions without inspection. These requirements led the industrial turbomachinery market to be generally conservative and sensitive to every single change through modern compressor development. The consequence for industrial compressor designs are, that these have made only moderate development steps during the last 50 years.

This paper deals with a novel hybrid axial flow compressor, which combines the advantages of an conventional industrial compressor, such as good operating range and efficiency, with the advantages of gas turbine compressors, mainly the higher power density resulting in a higher stage pressure ratio. Furthermore, the surge robustness of the novel compressor blading has been strongly improved. Starting from scratch, the development began with comprehensive matrix studies in all areas of the design, taking into account aerodynamics, mechanics, rotor dynamics and power density in order to ascertain the overall optimum for this new hybrid generation. State of the art CFD analysis has been intensively used to optimize the compressor blading as well as the flow behavior of inlet and exit for the specified requirements and different compressor control mechanisms.

The novel hybrid compressor is designed for a volume flow of 930 000 m3/h and allows a scaling from 100 000 up to 1 500 000 m3/h of air. To verify the design, a rig — downscaled by the factor of 3 — was tested. The rig was intensively instrumented with thermocouples and pressure probes, a torquemeter, strain gauges, tip-timing probes, and transient pressure transducers. Besides the measurement of blading performance, inlet and exit flange-to-flange instrumentation has been used to collect performance data under a variety of industrial operating conditions. The compressor behavior will be presented with a focus on aerodynamic aspects. The analytical and experimental data will be discussed in detail.
Towards a Highly Efficient Small Scale Turboshaft Engine: Part II — Aero-Mechanical Turbine Design and Optimization
Authors: Siller U. , Kröger G. , Moser T. , Hediger S.
Published: September 2014

Aero-mechanical design of the turbine section of a small scale turboshaft engine is presented in this paper. A single stage high-pressure turbine (HPT) and the power turbine stage (PT), have been designed by means of automated optimization. This study demonstrates how multi-disciplinary optimization can be used effectively in today’s industrial development cycles with respect to timeframe and computational resources. Both, the aerodynamic performance and the mechanical blade behavior were subject to the optimization in a very high dimensional design space expressed by well above 100 free design parameters for the annular duct and the bladings of two axial stages. In the first part, this paper describes the design task and constraints in order to meet the requested thermodynamic cycle performance and fabricational requirements. In the second part, the optimization strategy is explained with focus on geometry parameterization, simulation setups for flow and structural analysis and acceleration techniques for the optimization itself.

Finally, a very promising resulting design is reviewed in terms of a detailed aerodynamic and mechanical assessment and regarding to the overall engine concept. This work contributes to the development of a highly efficient, light-weight propulsion system powering, beside a wide range of other possible applications, for example small aerial systems such as helicopter drones.

The engine prototype is expected to be tested the first time in 2014.
Operation of a Burner Rig for Thermal Gradient Cycling of Thermal Barrier Coatings
Authors: Feist J. P. , Sollazzo P. Y. , Pilgrim C. , Nicholls J. R.
Published: September 2014

Thermal barrier coatings (TBC), in combination with sophisticated cooling systems are crucial for the operation of highly efficient gas turbines. New generations of coatings will need to show increased cycling capability as a future energy mix will contain a high proportion of renewable energy which will be subject to rapid changes in supply. This will require gas turbines to be on stand-by to fill shortages in power supply with short notice. Furthermore, higher operating temperatures are sought to improve the efficiency of the engine. It is, therefore, an aim of the industry to find a coating composition or structure which will enable the operation at temperatures greater than 1250°C and with high cycling capability.

Test methods are required to meet these new operating conditions to validate new coatings. The maximum temperature limit of commonly used isothermal or cyclic oxidation tests is usually the temperature at which the substrate will start to significantly oxidise. However, there is the technical need to test the ceramic top layer at elevated surface temperatures up to 1500°C while keeping the substrate ‘cool’. Such capability would allow the effects of ceramic sintering, and deposit induced damage to be assessed at the TBC surface. This only can be performed on a complete coating system, when a thermal gradient is established throughout the coating.

This paper reviews a burner test facility, designed and built by Sensor Coating Systems Ltd. (SCS), which combines severe and frequent cycling with the exposure of the coating to high surface temperatures and active cooling of the substrate. Further, this test can include thermal shock by active cooling of the surface at the end of each cycle. The paper will consider different operating conditions and will review experiences in building and operating the rig, including results from thermal barrier coating tests on electron beam physical vapour deposition (EBPVD) and atmospheric plasma spray (APS) samples. Further, the rig is capable of testing optical techniques such as pyrometry and thermographic phosphor thermometry for measuring surface temperature in controlled laboratory conditions and example of this will be presented. The paper also will reflect on the ISO 13123:201 standard for this type of test.
Numerical and Experimental Analysis of Radial Clearance Influence on Rotor and Stator Clocking Effect by Example of Model High Loaded Two Stage Compressor
Authors: Mileshin V. I. , Savin N. M. , Kozhemyako P. G. , Druzhinin Ya. M.
Published: September 2014

It is ascertained that optimization of bladed rows clocking positions is an effective tool for control of bladed rows unsteady interaction in a multistage turbo machine and could be used equally in refinement and in compressor design stages [1–3].

Up to now the clocking effect issue of highly loaded rotors in compressor stages was not investigated due to growing design complexity of the experimental facility. The issue concerning tip clearances value influence on rotor and stator clocking effects was not studied as well.

In the frames of this work a two stage compressor (HPC2 with the designed pressure ratio π* = 3.7 [4]) with a unique design was developed to investigate influence of tip clearance values on rotor and stator clocking effect. In order to realize the clocking effect the HPC2 compressor has the following parameters: numbers of stator blades are identical for all stators and equal to ZIGV = ZS1 = ZS2 = 68; number of R1 blades is two times less than of R2: ZR2 = 56, ZR1 = 28.

This work studies HPC2 compressor performances at 3 R1 and R2 tip clearances — dtip = 0.5mm – nominal clearance, 0.75mm, and 1.0 mm – increased clearance. Clocking effects of stators and rotors are tested at 0.5mm and 0.75mm tip clearances for two values of corrected rotational speeds — n = 0.7 and 0.88. As shown, variations in max. efficiency from maximum to minimum when changing the clocking position both the stator and the rotor reach Δη*ad≈0.008÷0.012 at dtip = 0.5mm or Δη*ad≈0.007÷0.008 at dtip = 0.75mm.

For more detailed analysis of the tip clearance influence on rotor and stator clocking effect a mathematical model of HPC2 was developed on base of through flow 3D viscous unsteady flow computations in the HPC2 compressor rows [3–4]. In full unsteady statement the calculation domain includes the following number of blade rows: IGV = 2, R1 = 1, S1 = 2, R2 = 2; S2 = 2.
Review of Platform Cooling Technology for High Pressure Turbine Blades
Authors: Wright Lesley M. , Malak Malak F. , Crites Daniel C. , Morris Mark C. , Yelavkar Vikram , Bilwani Rajesh
Published: September 2014

With the relatively large surface area of the platform of the gas turbine blades being exposed directly to the hot, mainstream gas, it is vital to efficiently cool this region of the blades. This region is particularly difficult to protect due to the strong secondary flows developed at the airfoil junction (formation of the leading edge horseshoe vortex) and circumferentially across the blade passage (strengthening passage vortex moving from the pressure side to the suction side of the passage). Over the past decade, researchers and engine designers have attempted to combat the enhanced heat transfer to the blade platform by implementing both frontside and backside novel cooling techniques. This paper presents a review of platform cooling technology ranging from frontside film cooling via stator-rotor purge flow, mid-passage purge flow, and discrete film holes to backside cooling achieved via impinging jet arrays or cooling channels. To gain a full understanding of state-of-the-art cooling technology, recent patents, journal articles, and conference proceedings are included in this review.
Active Tip Clearance Flow Control for an Axial-Flow Turbine Rotor Using Ring-Type Plasma Actuators
Authors: Matsunuma Takayuki , Segawa Takehiko
Published: September 2014

Tip leakage flow through the small gap between the blade tip and the casing wall in turbomachinery reduces the aerodynamic performance of the blade. New ring-type dielectric barrier discharge (DBD) plasma actuators have been developed to facilitate active control of the tip leakage flow of a turbine rotor. In the present study, the ring-type plasma actuators consisted of metallic wires coated with insulation material, mounted in an insulator embedded in the tip casing wall. For the fundamental experiments using a flat plate and a single airfoil with tip clearance, particle image velocimetry (PIV) was used to obtain two-dimensional velocity field measurements near the plate and blade tip regions. From flat plate experiments in a static flow field, it was confirmed that the operation of the plasma actuator generates an upward flow at the corner between the blade tip and the casing wall, and this forms a perpendicular obstacle to the tip leakage flow. In flat plate experiments on tip leakage flow in a wind tunnel, the forcibly-induced tip leakage flow was successfully dissipated by means of the plasma actuator flow control. In single airfoil experiments, the tip leakage flow was also reduced by the plasma actuator. In annular turbine rotor experiments, the plasma emission at the blade tip and its motion with blade rotation were determined. Single-element hot-wire anemometry was used to measure the turbulence intensity distributions at the turbine rotor exit. The amplitude of input voltage for the plasma actuator was varied from ±3.0 to ±6.0 kV. The high turbulence intensity region created by the tip leakage flow was reduced with an increase in the input voltage of the plasma actuator.
Influence of Ru, Mo and Ir on the Behavior of Ni-Based MCrAlY Coatings in High Temperature Oxidation
Authors: Yuan Kang , Peng Ru Lin , Li Xin-Hai , Johansson Sten , Wang Yan-Dong
Published: September 2014

To improve the oxidation/corrosion resistance of MCrAlY coatings (M for Ni and/or Co), elements like Y, Si and Ta have been added into the coatings in past decades. In this study the oxidation performance of a Ni-based MCrAlY coating with small proportion of Ru, Mo and/or Ir were investigated after high-temperature exposure. The oxidation tests were carried out at 900°C, 1000°C or 1100°C. The micro structure study showed that the addition of Ru, Mo and/or Ir had significant influence on the oxidation behavior at the coating surface and the microstructural evolution in the material. The microstructural evolution was quantitatively evaluated by measuring the phase degradation of β-NiAl in the coating and γ′-Ni3Al in the substrate of superalloy. Since no oxides of Ru, Mo and Ir were found on the coating surface, it was believed that the effects by those elements were mainly due to their dissolution in the metallic phases in the coatings.
Optimization of Film-Cooling Holes Shape by Principal Components Method and CFD
Authors: Vinogradov Kirill A. , Didenko Roman A. , Karelin Dmitry V. , Remizov Alexander E.
Published: September 2014

The development of effective cooling systems is one of the most actual and difficult problems of development of gas-turbine engines. Significant amount of works is devoted to the research of characteristics and optimization of the film cooling holes shape. Different shapes of holes are investigated at various working conditions. However, the development of optimization methods of film cooling holes shape is covered insufficiently.

Optimization methods are widely applied for solving modern optimization problems on the basis of surrogate models. The most famous one is the method of response surface. The aim of this method is to construct an approximation response surface of the objective function. It is also necessary to create and evaluate many variants of the investigation object design and therefore to conduct the same amount of CFD calculations or experiments in order to provide the acceptable design of an experiment (DOE). Such number of points on the response surface results in extreme computational costs increasing, i.e. so called “Curse of Dimension”.

The main objective of the current paper is to develop the method to improve film-cooling holes shape along with minimum time and computational costs and acceptable accuracy of the surrogate model. Besides, the analysis of optimization results is considered.

In the structural-parametrical analysis, methods of transformation of multidimensional information are used to generalize geometrical parameters of film cooling holes shape. The optimization process is carried out by means of random search methods. As a result, the “fast” surrogate model featuring acceptable accuracy was obtained. Then by regression analysis methods the structural coupling equation (SCE) between the principal components of holes shape and its cooling effectiveness was obtained. The authors carried out the study of the optimum variant of a film cooling holes shape in the space of the principal components. Geometrical information of the optimized variant has been obtained and further calculation of its characteristics was carried out by means of the “reverse recovery” technology developed.

Finally, the offered method enables us to get the film-cooling holes geometry variant with a higher level of cooling effectiveness compared to the considered classical variants. Thus, the offered way of optimization provides eightfold decrease in necessary number of CFD calculations compared to the classical approach (DOE and the response-surface methods).
A Step Further on the Control of Acceleration of Gas Turbines With Controlled Variable Geometry and Combustion Emissions
Authors: Barbosa João Roberto , Bringhenti Cleverson , Tomita Jesuíno Takachi
Published: September 2014

The paper aim is to study the transient performance using fuel flow schedule, variable geometry compressor control and combustion emissions for a simple turbojet engine in a thrust class of 5kN. This engine is under development, it was designed, manufactured and are being tested in test bench, it is composed by a 5-stage axial flow compressor, an annular combustor, an uncooled turbine and a convergent nozzle. The engine was originally designed to run on kerosene but other types of fuels, as biofuels, are intended to be used, having in mind a turboshaft in a class of 1.2 MW for power generation purpose. PID control is being studied to determine the appropriate setting of the VIGV in conjunction with a prescribed fuel flow injection necessary to accelerate, or decelerate, the engine from 80% to full thrust in a prescribed time interval. The engine control system is being studied during this engine design phase, so that all the components characteristics needed are being synthesized using in-house developed computer codes: compressor design and performance; combustion chamber design and performance; turbine design and performance; whole engine performance. The engine is required to accelerate from 80% to full thrust in a short time interval, which is also a limitation imposed to the control system. Compressor surge margin is controlled during accelerations using controlled positioning of the VIGV at each engine speed. The engine running lines, for accelerations and decelerations are shown and commented. They served as basis for the design of the engine control logic and hardware. The combustor was designed for kerosene, but other types of fuels can be burned, with the lower heating value and all the necessary parameters recalculated using reaction mechanisms, reactor network and stability loops approach.
Analysis of the Noise Level Generated by Axial Flow Fan Composed of Radial-Bladed Centrifugal Rotor
Authors: Borges S. S. , Barbieri R. , Zdanski P. S. B.
Published: September 2014

The objective of this work is to present, by means of experimental, analytical and numerical techniques that sound pressure level generated by radial-bladed centrifugal fans of electric motor cooling systems may be expressed by a logarithmical ratio of the peripheral velocity of rotor, volumetric flow and efficiency of the fan. The proposed methodology proved to be efficient and simple in the prediction of generated noise by radial-bladed centrifugal fans of TEFC motors with accuracy of ± 3 dB.

In addition, the acoustic resonance mode of the fan cavity were determined by means of numerical simulations, which its results were validated through experiments using waterfall spectrum.
Real-Time Burst Signal Removal Using Multi-Color Pyrometry Based Filter for Improved Jet Engine Control
Authors: Wang Guanghua , Estevadeordal Jordi , Harper Sean P. , Nirmalan Nirm
Published: September 2014

Online line-of-sight (LOS) pyrometer is used on certain jet engines for diagnosis and control functions such as hot-blade detection, high-temperature limiting, and condition-based monitoring. Hot particulate bursts generated from jet engine combustor at certain running conditions lead to intermittent high-voltage signal outputs from the LOS pyrometer which is ultimately used by the onboard Digital Engine Controller (DEC).

To study the nature of hot particulates and enable LOS pyrometer functioning under burst conditions, a Multi-Color Pyrometry (MCP) system was developed under DARPA funded program and tested on an aircraft jet engine. Soot particles generated as by-product of combustion under certain conditions was identified as the root cause for the signal burst in a previous study. The apparent emissivity was then used to remove burst signals.

In current study, the physics based filter with MCP algorithm using apparent emissivity was further extended to real-time engine control by removing burst signals at real time (1MHz) and at engine DEC data rate. Simulink models are used to simulate the performances of the filter designs under engine normal and burst conditions. The results are compared with current LOS pyrometer results and show great advantage. The proposed model enables new LOS pyrometer design for improved engine control over wide range of operating conditions.
Crossflows From Jet Array Impingement Cooling: Effects of Hole Array Spacing, Jet-to-Target Plate Distance, and Reynolds Number
Authors: Lee Junsik , Ren Zhong , Ligrani Phil , Fox Michael D. , Moon Hee-Koo
Published: September 2014

Data which illustrate the combined and separate effects of hole array spacing, jet-to-target plate distance, and Reynolds number on cross-flows, and the resulting heat transfer, for an impingement jet array are presented. The array of impinging jets are directed to one flat surface of a channel which is bounded on three sides. Considered are Reynolds numbers ranging from 8,000 to 50,000, jet-to-target plate distances of 1.5D, 3.0D, 5.0D, and 8.0D, and steamwise and spanwise hole spacing of 5D, 8D, and 12D, where D is the impingement hole diameter. In general, the cumulative accumulations of cross-flows, from sequential rows of jets, reduce the effectiveness of each individual jet (especially for jets at larger streamwise locations). The result is sequentially decreasing periodic Nusselt number variations with streamwise development, which generally become more significant as the Reynolds number increases, and as hole spacing decreases. In other situations, the impingement cross-flow results in locally augmented Nusselt numbers. Such variations most often occur at larger downstream locations, as jet interactions are more vigorous, and local magnitudes of mixing and turbulent transport are augmented. This occurs in channels at lower Reynolds numbers, where impingement jets are confined by smaller hole spacing, and smaller jet-to-target plate distance. The overall result is complex dependence of local, line-averaged, and spatially-averaged Nusselt numbers on hole array spacing, jet-to-target plate distance, and impingement jet Reynolds number. Of particular importance are the effects of these parameters on the coherence of the shear layers which form around the impingement jets, as well as on the Kelvin-Helmholtz instability vortices which develop within the shear interface around each impingement jet.
Turbofan Engine Health Assessment From Flight Data
Authors: Aretakis N. , Roumeliotis I. , Alexiou A. , Romesis C. , Mathioudakis K.
Published: September 2014

The paper presents the use of different approaches to engine health assessment using on-wing data obtained over a year from an engine of a commercial short-range aircraft. The on-wing measurements are analyzed with three different approaches, two of which employ two models of different quality. Initially, the measurements are used as the sole source of information and are post-processed utilizing a simple “model” (a table of corrected parameter values at different engine power levels) to obtain diagnostic information. Next, suitable engine models are built utilizing a semi-automated method which allows for quick and efficient creation of engine models adapted to specific data. Two engine models are created, one based on publicly available data and one adapted to engine specific on-wing “healthy” data. These models of different detail are used in a specific diagnostic process employing model-based diagnostic methods, namely the Probabilistic Neural Network (PNN) method and the Deterioration Tracking method. The results demonstrate the level of diagnostic information that can be obtained for this set of data from each approach (raw data, generic engine model or adapted to measurements engine model). A sub-system fault is correctly identified utilizing the diagnostic process combined with the engine specific model while the Deterioration Tracking method provides additional information about engine deterioration.
The Influence of Real Gas Effects on ICE-ORC Turbine Flow Fields
Authors: Zhang Lei , Zhuge Weilin , Zhang Yangjun , Peng Jie
Published: September 2014

This paper presents a quantitative comparison of the flow fields of a radial turbine between real gas and perfect gas models for the internal combustion engine (ICE) organic Rankine Cycle (ORC) application. Three-dimensional turbulent Navier-Stokes simulations are carried out using CFD code NUMECA FINE™/TURBO, which is linked to an accurate thermodynamic model for organic working fluid R123 in the form of thermodynamic tables. Four turbine operating conditions including the design point and three part-load points, the inlet compressibility factors of which are 0.82–0.89, are analyzed to discuss the differences of flow fields. Obvious derivations of thermodynamic parameters are investigated in the turbine flow fields. The derivations of speed of sound and density at the nozzle inlet are about 15–20%. There exist about 10m/s value differences in the nozzle outlet velocity evaluation, and furthermore a difference of 10 degrees in the rotor inlet incidence angle comparison. The derivations of relative Mach number are about 20–35% in the rotor outlet near the shroud surface. More than 30% differences are shown in the comparison of turbine total temperature drops. Other thermodynamic parameters show much smaller derivations. The differences of thermodynamic parameters lead to a 1–3% larger in mas flow rate, 1–2% larger in isentropic efficiency and 6–8% smaller in specific power comparison. However, there do not exist obvious differences on thermodynamic parameters distributions in the flow fields. The similar flow fields provide a suggestion that perfect gas model may be an acceptable model for turbine preliminary design and one-dimensional analysis in this gas thermodynamic region, and also the real gas flow fields simulated can be used as a start point to refine the turbine design.
Heat Transfer Performance of a Transonic Turbine Blade Passage in Presence of Leakage Flow Through Upstream Slot and Mateface Gap With Endwall Contouring
Authors: Roy Arnab , Jain Sakshi , Ekkad Srinath V. , Ng Wing F. , Lohaus Andrew S. , Crawford Michael E.
Published: September 2014

Comparison of heat transfer performance of a non-axisymmetric contoured endwall to a planar baseline endwall in presence of leakage flow through stator-rotor rim seal interface and mateface gap is reported in this paper. Heat transfer experiments were performed on a high turning (∼127°) turbine airfoil passage at Virginia Tech’s transonic blow down cascade facility under design conditions (exit isentropic Mach number 0.88 and 0° incidence) for two leakage flow configurations — 1) mateface blowing only, 2) simultaneous coolant injection from the upstream slot as well as mateface gap. Coolant to mainstream mass flow ratios (MFR) were 0.35% for mateface blowing only, whereas for combination blowing, a 1.0% MFR was chosen from upstream slot and 0.35% MFR from mateface. A common source of coolant supply to the upstream slot and mateface plenum made sure the coolant temperatures were identical at both upstream slot and mateface gap at the injection location. The contoured endwall geometry was generated to minimize secondary aerodynamic losses. Transient IR (Infrared) thermography technique was used to measure endwall surface temperature and a linear regression method was developed for simultaneous calculation of heat transfer coefficient (HTC) and adiabatic cooling effectiveness (ETA), assuming a 1D semi-infinite transient conduction. Results indicate reduction in local hot spot regions near suction side as well as area averaged HTC using the contoured endwall compared to baseline endwall for all coolant blowing cases. Contoured geometry also shows better coolant coverage profiles further along the passage. Detailed interpretation of the heat transfer results along with near endwall flow physics has also been discussed.
Investigation of a Compressor Rotor Non-Synchronous Vibration With and Without Fluid-Structure Interaction
Authors: Gan Jiaye , Im Hong-sik , Espinal Daniel , Lefebvre Alexis , Zha Ge-Cheng
Published: September 2014

This paper study the non-synchronous vibration (NSV) of a high speed multistage axial compressor using rigid blade and vibrating blade with fluid-structural interaction (FSI). The unsteady Reynolds-averaged Navier-Stokes (URANS) equations and mode based structural dynamic equations are solved. A low diffusion E-CUSP Reimann solver with a 3rd order WENO scheme for the inviscid fluxes and a 2nd order central differencing for the viscous terms are employed. A 1/7th annulus sector of IGV-rotor-stator is used with a time shifted phase lag BC at circumferential boundaries. An interpolation sliding boundary condition is used for the rotor-stator interaction. The URANS simulation for rigid blades shows that the leading edge (LE) tornado vortices, roughly above 80% rotor span, travel backwards relative to the rotor rotation and cause an excitation with the frequency agreeing with the measured NSV frequency. The predicted excitation frequency of the traveling vortices in the rigid blade simulation is a non-engine order frequency of 2603 Hz, which agrees very well with the NSV rig testing. For the FSI simulation, the results show that there exist two dominant frequencies in the spectrum of the blade vibration. The lower dominant frequency is close to the first bending mode. The higher dominant frequency close to the first torsional mode agrees very well with the measured NSV frequency. The simulation conducted in this paper appears to indicate that the NSV is excited by the traveling vortex.
Thermodynamic Analysis of a SCO2 Part-Flow Cycle Combined With an Organic Rankine Cycle With Liquefied Natural Gas as Heat Sink
Authors: Zhang Hanzhen , Shao Shuai , Zhao Hang , Feng Zhenping
Published: September 2014

The supercritical carbon dioxide (SCO2) part-flow cycle can achieve higher efficiency compared with conventional Brayton cycle as it can avoid the pinch point problem inside the regenerative heat exchanger. To recover the waste heat from the pre-cooler in the cycle and improve the overall cycle efficiency, a new integrated power system driven by nuclear reactor is proposed to achieve the energy cascade utilization. This system combines a SCO2 part-flow cycle with an organic Rankine cycle (ORC) using liquefied natural gas (LNG) as heat sink to utilize the cold energy of LNG. In this paper a mathematical model is established to simulate the SCO2 part-flow cycle coupled with an ORC under steady state condition, and a thermodynamic parametric analysis is conducted to investigate the effects of some key parameters, including the turbine inlet pressure, the turbine and compressor isentropic efficiency and pressure drop ratio, on the system performance. The results indicate that the integrated power system is effective to recover the waste heat and may achieve the overall cycle thermal efficiency of 52.12% under the operating conditions of 20MPa, 800K and part-flow ratio 0.68, which can be further improved with parametric optimization of the system.
A Practical Experiment and FEM Combined Method for Evaluation of Steam Turbine Diaphragm’s Creep Life
Authors: Meng C. , Xin J. H. , Ye C. , Han H.
Published: September 2014

The security and stability of the steam turbine equipments are of great importance for thermal power plants. The diaphragm of steam turbine is usually tested as lack of stiffness due to creep deformation in the practical operation. In this paper, a creep life assessment of the diaphragm of a practical steam turbine was presented. The finite element model of the diaphragm was construed, and its accuracy was verified via a defection test. The classic Norton creep law was applied to the calculation process and its creep parameters were determined via iteration to obtain an appropriate value which would make the calculated deformation data equal to the practical experimental data. The creep life of the diaphragm was obtained based on the appropriate finite element model and the right creep parameters. The result shows that the method is an effect way in the engineering field to evaluate the creep life of components of steam turbine and gas turbine’s equipment in the situation that the material property is unknown.
MAP: A CFD Package for Turbomachinery Flow Simulation and Aerodynamic Design Optimization
Authors: Ning Fangfei
Published: September 2014

This paper presents an in-house CFD package and its derivations for turbomachinery flow simulation as well as aerodynamic design optimization which have been employed in several primary Chinese aero-engine institutions. The package contains TurboMesh, a highly automated mesh generation code for turbomachinery, and MAP, a CFD solver for general purpose. Besides the programming strategies, the numerical schemes, and the parallelization methods adopted in the codes are outlined. Emphasis is placed on a novel mixing plane model used in MAP. The proposed model satisfies flux conservation property and very robust in actual usage. Additionally, further improved performance of the model can be achieved by applying a technique similar to the perfect matching layer for non-reflecting boundary conditions. On the basis of MAP, a set of derivations have also been developed. They include several versions of MAP which are based on specific flow models, respectively, an inverse code for the design optimization of 3D blade shape, an improved-delayed-detached-eddy-simulation based code, and a scale-adaptive-simulation based code. Some of these versions of MAP are briefly introduced and demonstrated through a few examples except for the inverse code in which a direct method proposed by the author is explained with a few more words. Illustrations show the applicability of the inverse code for the design of compressor blades in practical multi-blade row environment. By embedding MAP with an in-house numerical optimization package, the numerical optimization of the 2D/3D blade shape can be realized. Some examples for 3D aerodynamic optimizations of compressors are presented.
Experimental Investigations on Cooling Hole Diameters of an Impingement-Effusion Cooling System
Authors: Li Xiaopei , He Yuelong , Mao Ronghai , Fan Renyu , Xu Rongsun , Zhang Bo
Published: September 2014

Cooling holes are playing critical roles in the impingement-effusion cooling system. The present experimental investigation have studied the effects of the hole diameters on the cooling efficiency and discharge coefficient. The main flow was heated up via combustion, with a maximum outlet temperature 850K. Cooling air, on the other hand, is supplied by an MAM-200 compressor. Temperature distribution on the hot side of the effusion plate is measured by a thermal infrared imaging camera calibrated by 5 thermocouples embedded in the tested plate. A filter is used to minimize the radiation from the hot gas.

Cooling holes, both the effusion holes and impingement holes are designed in staggered arrangement; and the number of effusion holes is almost equal to the impingement holes. The inclination angle of the effusion holes in the streamwise direction is 20 degree with respect to the liner surface. The pressure drop across the cooling system ranges from 3% to 5% of the total pressure of the inlet flow. Temperature distributions are measured with an infrared thermal imaging system. The length of the holes is kept to be constant. Three various effusion hole diameters/length ratio 0.3, 0.35, 0.43, with impingement hole diameters/length ratio 0.4, 0.47 and 0.6 correspondingly, have been tested and compared. The experimental results indicate that with the increasing of cooling hole diameters, discharge coefficient is also increased. Nevertheless, cooling efficiency reached its maximum with the effusion hole diameter/length ratio 0.35.
CFD Investigation of the Influence of Inflow Conditions on an Aggressive Axial Exhaust Diffuser for Steam Turbines
Authors: Xu Lei , Hedlund Lars , Nilsson Åsa
Published: September 2014

The aerodynamic performance of an aggressive axial steam turbine exhaust diffuser is investigated. CFD simulations are carried out with the last stage rotor included. As expected, results reveal that the last stage tip leakage has a significant influence on the performance of the diffuser. A comparison is made with results from calculations of a conservative diffuser. This shows that although the high total pressure in the tip leakage jet helps to prevent casing separation, it can ruin the recovery if the diffusion is too low. Then the high velocity at the casing can reach the outlet of the diffuser, resulting in a more non-uniform exit flow distribution. This in turn reduces the pressure recovery. This effect is of special importance when the non-dimensional length is short, as for the investigated diffusers.
Combined Low and High Cycle Fatigue Tests on Full Scale Turbine Blades
Authors: Chen Xia , Yan Xiaojun
Published: September 2014

Since the unavoidable vibrations wildly exist in the operating environment of turbine blades, the assessment of the combined low and high cycle fatigue (CCF) properties are of paramount importance for turbine blades. To study turbine blades’ CCF properties accurately, one important and useful way is to carry out the CCF tests on full scale turbine blades. When conducting the CCF tests of turbine blades, there are two challenges. One is finding ways/paths to transfer the low cycle load and the high cycle load properly, and the other is the determination of high cycle vibration stress level in the working condition. In this investigation, a new experimental method, in which full scale turbine blades are loaded by a special design test system providing high cycle loads and low cycle loads without interference, is proposed to conduct CCF life tests, while an analysis method (reverse method) based on the CCF tests was developed to estimate the high cycle vibration stress level. The design of the test system and the principle of the reverse method are fully explained in this paper. Test results show that the test system can apply HCF/LCF combined loading on the key section of the full scale turbine blade, and successfully simulated the working state. Meanwhile, the working condition vibration stress range of the blade is obtained through the reverse method.
An Adjoint-Based Optimization Method for Constrained Aerodynamic Shape Design of Three-Dimensional Blades in Multi-Row Turbomachinery Configurations
Authors: Walther Benjamin , Nadarajah Siva
Published: September 2014

This paper develops the discrete adjoint equations for a turbomachinery RANS solver and proposes a framework for fully-automatic gradient-based constrained aerodynamic shape optimization in a multistage turbomachinery environment. The systematic approach for the development of the discrete adjoint solver is discussed. Special emphasis is put on the development of the turbomachinery specific features of the adjoint solver, i.e. on the derivation of flow-consistent adjoint inlet/outlet boundary conditions and, to allow for a concurrent rotor/stator optimization and stage coupling, on the development of an exact adjoint counterpart to the non-reflective, conservative mixing-plane formulation used in the flow solver. The adjoint solver is validated by comparing its sensitivities with finite difference gradients obtained from the flow solver. A sequential quadratic programming algorithm is utilized to determine an improved blade shape based on the objective function gradient provided by the adjoint solution. The functionality of the proposed optimization method is demonstrated by the redesign of a single-stage transonic compressor. The objective is to maximize the isentropic efficiency while constraining the mass flow rate and the total pressure ratio.
CFD Analysis of Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade
Authors: Alameldin Ahmad , El-Gabry Lamyaa A. , Fridh Jens , Saha Ranjan
Published: September 2014

Operating at temperatures well above their melting point, gas turbines’ components are subject to terribly high thermal stresses. In order to keep them intact and performing, different cooling techniques are implemented. One of these methods is film cooling.

Film cooling implementation in vane cascades has a potential loss expense. Proper assessment of its impact on the vane performance has to be conducted. The CFD approach of modeling each hole and cooling tube autonomously is very computationally expensive. In the current work an assessment of a new, more computationally efficient CFD approach for modelling film cooling was conducted on a vane cascade operating in the transonic regime (M = 0.89). The film cooling holes were represented by orifice boundary condition at the vane surface, omitting the need to model internal coolant plenum and cooling tubes mesh, resulting in 180% reduction in grid size and attributed computational cost interpreted in 300% saving in computation time.

Uncooled, and film cooled with different configurations and at different blowing ratios (BR) simulations were performed and compared to experimental measurements. A good agreement was obtained for the exit flow angles, vorticity and aerodynamic loss for all the cases (uncooled and cooled). Pitch-averaged exit flow angle outside endwalls regions remains unchanged for all cooling configurations and blowing ratios. The aerodynamic loss was found to be more sensitive to increasing the blowing ratio on the suction side than on the pressure side.

The proposed approach of coolant injection modeling is shown to yield reliable results, within the uncertainty of the measurements in most cases. Along with lower computational cost compared to conventional film cooling modeling approach, the new approach is recommended for further analysis for aero and thermal vane cascade flows.
Parametric Optimization of Zeotropic Mixtures Used in Low-Temperature Organic Rankine Cycle for Power Generation
Authors: Xi H. , Li M. J. , He Y. L. , Yang W. W. , Li Y. S.
Published: September 2014

In the design and optimization of the ORC system, the selection of working fluid is one of the most important factors that should be considered. In this work, considering different heat sources with their temperatures ranging from 80 to 120 °C, 8 different zeotropic mixtures were proposed and their thermodynamic and economic performance for two types of traditional ORC systems (i.e. basic organic Rankine cycle, BORC and organic Rankine cycle with internal heat exchanger, IHORC) were investigated. Firstly, economic analysis were performed for both systems; Secondly, genetic algorithm (GA) was then introduced to determine the optimal fractions and other operation parameters for zeotropic mixtures under different working conditions and systems, the algorithm implementation process was described. Thirdly, the optimization studies were performed by using annual cash flow as the objective function. The optimal thermodynamic performance of different zeotropic mixtures and their components were both calculated and compared. For the different heat sources temperatures, the optimal zeotropic mixtures and their optimal fraction were recommended according to the calculated results.
A Matching Method for Two-Stage Turbocharging System
Authors: Liu Yanbin , Zhuge Weilin , Zhang Yangjun , Zhang Shuyong , Zhang Junyue , Huo Xuemin
Published: September 2014

The turbine system of a two-stage turbocharger composed of high pressure turbine, low pressure turbine and by-pass valve decides distribution and utilization of exhaust gas energy and influence performance of two-stage turbocharger in whole operational conditions. Besides, characteristics of turbine is expressed by envelop line of characteristic lines in different speeds. So turbine can be conveniently selected compared with compressor with similarity theory. Therefore two-stage turbocharger matching begins from turbine system matching in the paper. In two-stage turbocharger, cooler efficiency, cooler loss and by-pass valve open besides turbochargers will influence turbocharging system performance and design of cooler and by-pass valve are important contents of turbocharging system matching. The paper matched inter cooler, by-pass valve open, compressors and turbines jointly. Calculation model for turbocharger matching was built, and turbine performance is get from reference turbine based on similarity theory; influence of compressor ratio distribution, cooler efficiency and pressure drop in cooler imposing on compressor work was analyzed; and influence of turbine flow capacity and by-pass valve imposing on output working in expanding process was studied; the method for matching of two-stage turbocharging system in whole operational condition is studied Matching analysis was made aiming at two-stage turbocharging system of a type of high power density diesel engine, and design for turbocharging system was finished. Matching result using the method is compared to matching result using traditional method. Analysis result proves that using the method matching points in different operational conditions are located in more reasonable zone of compressor MAP.
Hot-to-Cold CAD Geometry Transformation of Aero Engine Parts Based on B-Spline Morphing
Authors: Gaun Lilia , Bestle Dieter , Huppertz André
Published: September 2014

Typically, complex components are designed with a CAD system, where free-form surfaces are usually described by B-Spline or NURBS surfaces. Especially in turbo machinery, design is performed for hot working conditions. However, for manufacturing a cold, unloaded CAD model is required. A deformed component with smooth surfaces may also be needed e.g. for flow calculations at different operating points. Engine-part deformations caused by loads and temperature changes are calculated by a finite element program which produces a displacement field as output. In such cases the original B-Spline description may be used and the shape of the relevant surfaces may be adjusted via displacement of the B-Spline control-points. The procedure is to decompose the solid into part-surfaces, identify corresponding FE nodes, apply the FE displacement field to the projected points of the FE nodes, and determine displacements of the control points by solving linear fitting problems. The required position and tangential continuity can be achieved by special treatment of the edge curves and their neighboring control points. Treatment of singularities of the linear fitting problem at trimmed surfaces and inadequate FE meshes requires special focus. The presented generic morphing method was implemented as an extension to the CAD system Unigraphics NX 7.5 and can now be used as the basis for automated hot-to-cold geometry transformations. Applications to a compressor aerofoil and an exhaust mixer segment demonstrate the robustness and accuracy of the method.
Predictability of Rotating Stall and Surge in a Centrifugal Compressor Stage With Dynamic Simulations
Authors: González Díez Néstor , Smeulers Jan P. M. , Tapinassi Libero , Del Greco Alberto Scotti , Toni Lorenzo
Published: September 2014

A testing campaign performed on a centrifugal compressor test-loop showed that multiple areas of stable operation existed that were not expected. It was shown during the tests that after a first region of instability at low flow rates, a second stable region would appear at even lower flow rates. Dynamic simulations based on a modified Greitzer’s model have been performed in order to assess the predictability of these unsteady phenomena. The study is focused on matching the dynamic simulations and the experimental results, concerning the different stable and unstable regions. It was of particular interest to verify the simulation results for the first region of instability, where rotating stall is suspected to occur. In this region the time-averaged performance curve showed a positive, i.e. unstable, slope. The performance curve presents then two regions with negative, stable, slope. The model shows good agreement with the measured pressure fluctuations when operated in surge conditions, both in terms of fundamental surge cycle frequency and pulsation amplitude. Capturing the pressure variations for the rotating stall range proved more difficult and subject to the implemented performance curve, which is considered uncertain in this range. By tuning the performance curve in the corresponding range of flow rates, a satisfactory match with the test results could be achieved. The second range of stability was correctly captured by also modifying the compressor performance curve.