The ever increasing performance requirements of modern aero-engines necessitate the development of effective ways to improve efficiency and reduce losses. Casing temperature control is particularly critical from this point of view, since thermal expansion directly affects the blade tip clearance and thus the associated leakages. To limit the turbine tip flows, Active Clearance Control (ACC) systems have been implemented over the last decades. These systems are usually based upon impingement cooling, generated by a series of perforated manifolds enclosing the turbine casing. When dealing with aeroengine low pressure turbines, the current trend in increasing the engine by-pass ratio, so as to enhance the system propulsive efficiency, pushes the limits of ACC traditional design performance. The reduction of the pressure head at the ACC system inlet requires lower nozzle-to-target distances as well as denser impingement arrays to compensate the reduction of the jets’ Reynolds number. Literature correlations for the impingement heat transfer coefficient estimation are then out of their confidence range and also RANS numerical approaches appear not suitable for future ACC designs. In this work, methodologies for the development of accurate and reliable tools to determine the heat transfer characteristics of low pressure ACC systems are presented. More precisely, this paper describes a custom designed finite difference procedure capable of solving the inverse conduction problem on the target plate of a test sample. The methodology was successfully applied to an experimental setup for the measurement of the thermal loads on a target plate of a representative low pressure ACC impinging system. The experimental data is then used to validate a suitable numerical approach. Results show that RANS model is not able to mimic the experimental trends, while scale-resolving turbulence models provide a good reconstruction of the experimental evidences, thus allowing to obtain a correct interpretation of flow and thermal phenomena for ACC systems.

In real gas turbine engines, a gap/step interface commonly exits between upstream of the inlet guide vane endwall and combustor, called upstream endwall misalignment, due to the errors of assembly and the thermal expansion. This endwall misalignment, commonly being presented as the gap/step geometry with different heights, has a significant effect on the endwall heat transfer and film cooling coverage distributions. This paper presents a detailed experimental and numerical study on the effects of upstream endwall misalignment (step geometry) on the vane endwall heat transfer and film cooling in a transonic linear turbine vane passage. The experiment measurements were performed in a blowdown wind tunnel at simulated realistic gas turbine operating conditions (high inlet freestream turbulence level of 16%, exit Mach number of 0.85 and exit Reynolds number of 1.7 × 10 6 . Three types of upstream step geometry were tested at design blowing ratio (BR = 2.5) for the same vane profile: I) baseline geometry with zero-step height of Δ H = 0 mm; II) forward-facing step geometry with negative step height of Δ H = −5 mm; III) backward-facing step geometry with positive step height of Δ H = 5 mm. The endwall thermal load and film cooling coverage distributions were measured using transient infrared thermography, being presented as endwall Nusselt number Nu and adiabatic film cooling effectiveness η, respectively. Detailed comparisons of experiment measurements with numerical predictions were also presented and discussed for three types of upstream step configurations with Δ H = −5, 0, 5 mm, respectively. The numerical simulations were performed by solving the steady-state Reynolds Averaged Navier Stokes (RANS) with Realizable k-ε turbulence model, based on the commercial CFD solver ANSYS Fluent v.15. The effects of upstream step geometry were numerically studied, at the same design blowing ratio BR = 2.5, by solving the endwall Nusselt number, film cooling effectiveness and secondary flow field for various upstream step heights: three forward-facing step heights (from −8 mm to −3 mm), a baseline step height (0 mm), and four backward-facing step heights (from 3 mm to 10 mm). The results show the upstream forward-facing step geometry is beneficial for the endwall thermal load and film cooling, though the improvement is weak for all step heights (less than 10% decrease in endwall heat transfer and less than 10% increase in endwall film cooling). However, the upstream backward-facing step geometry is pernicious for the endwall heat transfer and film cooling, and the influence increases with the increasing upstream backward-facing step height. The backward-facing step geometry obviously alters near endwall flow field, leading to an enhancement (up to 20%) in endwall heat transfer and significant reduction (up to 60%) in endwall film cooling effectiveness.

There is a drive towards a broader range of fuels in industrial gas turbines, with higher levels of sulphur and potentially hydrogen. Due to these harsher environments, there is also a drive for corrosion resistant alloys and coatings. A number of key corrosion resistant superalloys, which are being employed to cope with these evolving conditions, exhibit primary creep. It is therefore imperative that fundamental material models, such as those for creep deformation, are developed to ensure they can accurately predict the material response to evolving operating conditions. The requirements for a creep model are complex. The model must be able to: predict forward creep deformation in regions dominated by primary loads (such as pressure); predict stress relaxation in regions dominated by secondary loads (such as differential thermal expansion); predict the effects of different creep hardening mechanisms. It is also clear that there is an interaction between fatigue and creep. With flexible operation, this interaction can be significant and should be catered for in lifing methods. A model that has the potential to account for the effect of plasticity on creep, and creep on plasticity is therefore desirable. In previous work the authors presented the concept for a backstress model to predict creep strain rates in superalloys. This model was fitted to a limited dataset at a single temperature. The approach was validated using simple creep-dwell tests at the same temperature. This paper expands on the previous work in several ways: 1) The creep model has been fitted over a wide range of temperatures. Including the effect of temperature in complex creep models presents a number of difficulties in model fitting and these are explored. 2) The model was fitted to constant load (forward creep) and constant strain (stress relaxation) tests since any creep model should be able to predict both forms of creep deformation. However, these are often considered separately due to the difficulty of fitting models to two different datasets. 3) The creep deformation model was validated on stress change tests to ensure the creep deformation response can cope with changes in response variables. 4) The approach was validated using creep-fatigue tests to show that the creep deformation model, in addition to our established fatigue models, can predict damage in materials under complex loading.

Active magnetic bearings (AMBs) are being increasingly utilized in industrial applications due to their advantages over conventional bearings. They offer very low friction and wear, variable stiffness and damping, and greater tolerance of rotating mass imbalance. These unique features of AMBs have enabled design of robust rotating machinery at much higher speed with higher power concentration. The present work discusses the design of a high temperature magnetic bearing for operation at an axial thrust load of 4448N, speed 20000 rpm and temperature 538°C. Various disk profiles were considered to lower peak stresses due to centrifugal forces, including uniform (rectangular), linear tapered and hyperbolic. The predictions showed that the hyperbolic profile reduced stresses by 60% compared to the rectangular profile enabling rotor disks to operate at much higher speed. A test bearing was built with the hyperbolic disc profile. An Iron-Cobalt alloy, commercially known as Hyperco 27 was utilized for the thrust disc for its high yield strength 570MPa, high saturation flux density of 2.35T and high resistivity of 250μΩ-mm. Hyperco 50A was selected for the bearing stator, due to the lower load requirement and cost. Magnetic circuit design assumptions for the axial thrust AMB included (1) relative permeability of the magnetic material was nearly infinite, (2) fringing at gap edges as well as leakage flux were negligible, and (3) the field within the circuit was homogeneous. The initial circuit design was improved using finite element magnetic field analysis. The effective force acting on the hyperbolic rotor determined the required number of turns and current for the electromagnetic coils. Extensive structural finite element analyses suggested not to use an interference fit of the attached disk with the shaft. Rather, it was decided to utilize a sleeve and lock-nut mechanism. Inconel 718 was used for the shaft due to its slightly higher thermal expansion coefficient than Hyperco 27. The thrust AMB containment vessel included thermally-insulated radial and axial adjustment bolts to position and align the AMB inside the vessel. The AMB rotating assembly was spun using an electric motor. The magnetic force generated by the AMB at room temperature was similar to its predicted value, with a 0.85 derating factor. The magnetic force was temperature dependent and was reduced to 65% of its room temperature value, at 538°C. The maximum operating speed reached thus far in the experimental study was 5000 rpm. The magnetic bearing force was nearly invariant with rotational speed at any given temperature (e.g., room and high), while the electric current was held constant. The design indicates that the novel magnetic thrust bearing should perform well at the target operating conditions of 4448N axial load at 538°C (1000 lb-f at 1000°F), and 20,000 rpm. This has been achieved thus far only up to 5000 rpm. The force appears to be very insensitive to motion induced eddy currents up to the present maximum speed of 5000 rpm. Future work will focus on reaching the full speed target of 20,000 rpm at 538°C and 4,448 N loading.

Better life assessment of hot-components of an aero-engine can help improve its reliability and service life, while, reducing associated maintenance cost. Accurate prediction of Thermo-Mechanical Fatigue (TMF) is one of the crucial aspects of life prediction. Therefore, fully resolved simulation methodologies have gained attention as an ingredient for solving TMF problems owing to their potential for providing comprehensive insights into a system having hot components undergoing transient loading during operation. The present work focuses on a multi-physics simulation-based approach for the life-prediction of a representative gas-turbine combustor liner with an objective of providing a complete framework for TMF analysis of an actual aero-engine combustor liner. The presented methodology consists of a coupling between Computational Fluid Dynamics (CFD) and Finite Element Method (FEM). Thermal loads on the representative aero-engine combustor are predicted using Conjugate Heat Transfer (CHT) modeling in the CFD analyses for different operating conditions suitable for a flight cycle. A load cycle is then constructed using these thermal loads and is transferred to the structural analysis to evaluate the stresses in the liner. Results are obtained regarding spatially varying thermal expansion resulting in inelastic strains as governed by temperature and rate dependent material behavior. Stress and plastic strain history information from the structural analysis are processed to predict the life of different regions of the combustor liner. Different simulation methods for conjugate heat-transfer, load-cycle, material property extraction, thermal-stresses, and fatigue are evaluated, and an overall methodology involving accuracy and reasonable computational cost is proposed. The proposed methodology is numerically verified, and the verification results are presented in this work.

This study is concerned with the film cooling effectiveness of the flow issuing from the gap between the NGV and the transition duct on the NGV endwall, i.e. the purge slot. Different slot widths, positions and injection angles were examined in order to represent changes due to thermal expansion as well as design modifications. Apart from these geometric variations, different blowing ratios (BR) and density ratios (DR) were realized to investigate the effects of the interaction between secondary flow and film cooling effectiveness. The experimental tests were performed in a linear scale-1 cascade equipped with four highly loaded turbine vanes at the Institute of Fluid Mechanics and Fluid Machinery of the University of Kaiserslautern. The mainstream flow parameters were, with a Reynolds number of 300,000 and a Mach number (outlet) of 0.6, set to meet real engine conditions. By using various flow conditioners, periodic flow was obtained in the region of interest (ROI). The adiabatic film cooling effectiveness was determined by using the Pressure Sensitive Paint (PSP) technique. In this context, nitrogen and carbon dioxide were used as tracer gases realizing two different density ratios DR = 1.0 and 1.6. The investigation was conducted for a broad range of blowing ratios with 0.25 ≤ BR ≤ 1.50. In combination with 10 geometry variations and the aforementioned blowing and density ratio variations 100 single operating points were investigated. For a better understanding of the coolant distribution, the secondary flows on the endwall were visualized by oil dye. The measurement results will be discussed based on the areal distribution of film cooling effectiveness, its lateral spanwise as well as its area average. The results will provide a better insight into various parametric effects of gap variations on turbine vane endwall film cooling performance — notably under realistic engine conditions.

Current and future designs for advanced turbine systems, such as Integrated Gasification Combined Cycle (IGCC), advanced Natural Gas Combined Cycle (NGCC), and the emerging supercritical CO 2 (SCO 2 ) systems require increasing turbine inlet temperature (TIT), which is well beyond the substrate melting temperature. The well-known approach is coating the turbine blade with thermal barrier coatings (TBC) combined with internal cooling channel in the substrate. However, due to thermally grown oxide (TGO) and thermal expansion mismatch stresses, TBC spallation failure is a major concern. Furthermore, neither the ceramic coating layer nor the metallic bond coat in current TBC system can provide structural support to house the internal cooling channels. In this research, a method to fabricate high temperature protective structural coating on top of critical gas turbine components by additive manufacturing (AM) technique using oxide dispersion strengthening (ODS) metal powder is presented. A novel combined mechanochemical bonding (MCB) plus ball milling process is utilized to produce near spherical and uniformly alloyed ODS powders. AM-processed ODS coating by direct energy deposition (DED) method on MAR-247 substrate, with laser powers from 100W to 200W were carried out. The ODS coated samples were then subjected to thermal cyclic loadings for over 2200 cycles. For comparison, in our earlier studies, under the same cyclic testing condition, typical tested TBC coupons showed spallation failure after ∼400 cycles. Correlation of the measured ODS coating Young’s modulus using a unique non-destructive micro-indentation testing method with evolution of the ODS microstructures are studied to identify optimum AM processing parameters for best performance of the ODS samples. In particular, stability of secondary γ′ phase in the ODS coating after thermal cycles is analyzed. Test results revealed a thin steady durable alpha alumina oxide layer on the best performance ODS samples. After 2,200 thermal cycles, strong bonding at ODS/substrate interface is also maintained for most of the ODS coated samples. Test results also showed stable substrate microstructure due to the protective ODS coating even after 2,200 thermal cycles. These preliminary test results showed strong potential for applications of AM-assisted ODS coating on advanced gas turbine components.

Recently, high capacity and high efficiency turbo machines such as turbo blowers and turbo compressors have been being developed. To keep up with this trend, magnetic bearings are increasingly being applied to turbo machines instead of air foil bearings. In this study, a hybrid magnetic bearing composed of a permanent magnet and an electromagnet was applied to 300-HP (horsepower) turbo compressors with rated speeds of 50,000 rpm. The length of the shaft should be reduced as short as possible so that the compressor rotor can rotate stably while maintaining a small vibration at a rotation speed of 50,000 rpm. In this study, the additional shaft length for the axial gap sensor is eliminated by applying a new layout in which the axial gap sensor is placed on the inner surface of the thrust magnetic bearing. No-load tests and full-load tests in the manufactured turbo compressors were performed to evaluate the performance of the designed magnetic bearings. The biggest obstacle to this development is the heat problems caused by compressor impellers and high-speed motors. Thanks to the cooling system using water and air, the problem of thermal expansion of the rotor in the axial direction can be avoided. Thrust forces were estimated using the control current for thrust magnetic bearings. The experiment is also performed to evaluate the vibration of each turbo compressor and the results are presented. Based on the evaluated vibration, it has been confirmed that the hybrid magnetic bearings and the new axial gap sensor layout can support the turbo compressors stably.

Because of the superior sealing characteristics compared to labyrinth seals, brush seals found an increased spread in turbomachinery in recent years. Their outstanding sealing performance results mainly from their flexibility. Thus, a very small gap between the rotor and bristle package can be obtained without running the risk of severe detrimental deterioration in case of rubbing. Rubbing between rotor and seal during operation might occur as a result of e.g. an unequal thermal expansion of the rotor and stator or a rotor elongation due to centrifugal forces or manoeuvre forces. Thanks to the flexible structure of the brush seal the contact forces during a rubbing event are reduced, however the frictional heat input can still be considerable. Particularly in aircraft engines with their thin and lightweight rotor structures the permissible material stresses can easily be exceeded by an increased heat input and thus harm the engine’s integrity. The geometry of the seal has a decisive influence on the resulting contact forces and consequently the heat input. The complex interactions between the geometric parameters of the seal and the heat input and leakage characteristics are not yet fully understood. This paper presents the investigation of the influence of the geometric parameters of a brush seal on the heat input into the rotor and the leakage behaviour. Two seals with different packing densities were tested under relevant engine conditions with pressure differences ranging from 1 to 7 bar, relative surface speeds ranging from 30 to 180 m/s and radial overlaps ranging from 0.1 to 0.4 mm. The transient temperature rise during the rub event was recorded with 24 thermocouples in close proximity to the rub contact embedded in the rotor structure. By comparing the temperature curves with the results of a thermal finite element analysis of the rotor the heat input into the rotor was calculated iteratively. It could be shown that the packing density has a decisive influence on the overall operating behaviour of a brush seal. Furthermore, results are obtained for the heat flux distribution between seal and rotor are shown.

The catcher bearing is a crucial part of the magnetic bearing system. It can support the rotor when the magnetic bearing is shut down or malfunctioning and limit the rotor’s position when large vibration occurs. The sleeve bearing has the advantages of a relatively large contact surface area, simple structure and an easily replaced surface. There are already many applications of the sleeve type catcher bearings in the industrial machinery supported by the magnetic bearings. Few papers though provide thorough investigations into the dynamic and thermal responses of the sleeve bearing in the role of a catcher bearing. This paper develops a coupled elastic deformation — heat transfer finite element (FEM) model of the sleeve bearing acting as a catcher bearing. The FEM model investigates the dynamic and thermal behavior when a flexible rotor drops onto the sleeve catcher bearing. The thermal load caused by the thermal expansion is also considered. The flexible rotor is composed of Timoshenko beam elements. A coulomb friction model is used to model the friction force between the rotor and the sleeve bearing surface. The contact force and 2-D temperature distribution of the sleeve bearing are obtained by numerical integration. To validate the FEM code developed by the author, firstly, both the mechanical and thermal static analysis results of the sleeve bearing model are compared with the results calculated by the commercial software, “SolidWorks Simulation”. Secondly, the transient analysis numerical results are compared with the rotor drop test results in reference 13. Additionally, this paper explores the influences of different surface lubrication conditions, different materials, such as stainless steel, bronze, and aluminum, on rotor-sleeve bearing’s dynamic and thermal behavior. This paper lays the foundation of the fatigue life calculation of the sleeve bearing and provides the guideline for the sleeve type catcher bearing design.

This work presents experimental and numerical investigations into the vibrations of turbocharger rotors on full-floating ring bearings with a circumferential oil-groove. The pressure distribution in the fluid-film bearings is calculated through the Reynolds equation using a highly efficient global Galerkin approach with suitable trial and test functions. The numerical efficiency of the method is markedly increased as the resultant linear system is solved symbolically, establishing a semi-analytical solution. The temperature in the oil-film may increase due to the mechanical power dissipation, affecting the pressure distribution and the load capacity of the bearing. Therefore, a reduced thermal energy model is implemented together with the Reynolds equation to account for the variable oil-viscosity and for the thermal expansion of the surrounding solids. The thermal energy balance equations are implemented in a transient form, i.e. including the time dependent temperature term. The corresponding system of nonlinear differential equations is efficiently solved, leading to a further significant reduction in simulation times. The hydrodynamic bearing model including the thermal effects is finally coupled with the equations of motion of a turbocharger rotor and numerical run-up simulations are compared with experimental results. The comparisons show that the numerical model captures adequately the dynamics of the system, giving precise information about the frequencies and the amplitudes of the synchronous and the self-excited subsynchronous rotor vibrations.

Today’s power market asks for highly efficient turbines which can operate at a maximum flexibility, achieving a high lifetime and all of this on competitive product costs. In order to increase the plant cycle efficiency, during the past years, nominal steam temperatures and pressures have been continuously increased. To fulfill the lifetime requirements and still achieve the product cost requirements, accurate mechanical integrity based assessments on cyclic lifetime became more and more important. For this reason, precise boundary conditions in terms of local temperatures as well as heat transfer coefficients are essential. In order to gain such information and understand the flow physics behind them, more and more complex thermal modelling approaches are necessary, like Computational Fluid Dynamics (CFD) or even Conjugate Heat Transfer (CHT). A proper application of calculation rules and methods is crucial regarding the determination of thermal stresses, thermal expansion, lifetime or creep. The aim is to exploit during turbine developments the limits of the designs with the selected materials. A huge effort especially in validation and understanding of those methodologies was done with detailed numerical investigations associated to extensive measurement studies at onsite turbines in operation. This paper focuses on the validation of numerical models based on CHT calculations against experimental data of a large intermediate pressure steam turbine module regarding the temperature distribution at the inner and outer casing for nominal load as well as transient shut-down.

Ceramic recuperators could enable microturbines to achieve higher fuel efficiency and specific power. Challenges include finding a suitable ceramic fabrication process, minimizing stray heat transfer and gas leakage, mitigating thermal stress, and joining the ceramic parts to neighboring metal components. This paper describes engine and recuperator design concepts intended to address these obstacles. The engine is sized to produce twelve kilowatts of shaft power, and it has a reverse-flow compressor and turbine. Motivations for this layout are to balance axial thrust forces on the rotor assembly; to minimize gas leakage along the rotating shaft; to reduce heat transfer to the compressor diffuser; to enable the use of a simple, single-can combustor; and to provide room for lightweight ceramic insulation surrounding all hot section components. The recuperator is an annular, radial counterflow heat exchanger with the can combustor at the center. It is assembled from segmented wafers made by ceramic injection molding (CIM). These are housed in a pressure vessel to load the walls mainly in compression, and are joined together by flexible adhesives in the cool areas to accommodate thermal expansion. A representative wafer stack was built by laser-cutting, laminating, and sintering tapecast ceramic material. The prototype was tested at temperatures up to 675°C, and the results were used to validate analytical and computational fluid dynamics (CFD) models, which were then used to estimate the effectiveness of the actual design. Turbomachinery efficiencies were also calculated using CFD, and allowances were made for additional losses like bearing friction and gas leakage. Based on these component performance estimates, a cycle model indicates the engine could achieve a net fuel-to-electrical efficiency of 21%, at a core weight including the recuperator of 11 kg, or about 1 kg/kW electric output.

What make the labyrinth seal in a compressor stator well different from the normal labyrinth seal are the inlet and outlet rotor-stator disc cavities. Due to the presence of rotating disc cavities, the windage heating and the swirl development are remarkable, which can have a great influence on the leakage characteristic. Besides, when compressor operates at different speeds, the rotor and stator grow differently owing to centrifugal expansion and thermal expansion. Hence the tip clearance which determines the leakage mass flow changes with the varying of rotational speed and temperature in the stator well. A rotating test rig with rotational speed 8100rpm and pressure ratio range 1.05∼1.3 was designed for the test of labyrinth seal in a compressor stator well. A cantilevered structure was used to entirely collect the mass flow for an accurate measurement. To know the working tip clearance precisely, the set up tip clearance was measured with plug gauges, while the radial displacements of rotating disc and stationary casing were measured separately with two high precision laser distance sensors. The total temperatures of airflow in the stator well were measured with thermocouples to analyze the proportion of windage heating among the inlet rotating disc cavity, outlet rotating disc cavity, and labyrinth seal segment. The disc and stator casing were manufactured with non-metallic materials to reduce heat dissipation. Furthermore, the circumferential velocity of the leakage flow was measured using probes to reveal the swirl development. Two-dimensional, axisymmetric swirl flow numerical simulations were carried out to provide insight into the flow field details, total temperature variation and swirl flow development in the stator well. The numerical results of discharge coefficient, windage heating and swirl ratio were compared with the experimental data. Of particular note is, the tip clearance of numerical model at a specific rotating speed was set to be the same with the actual working clearance which was measured in the experiment. The inlet and outlet parameters corresponded with the experimental conditions also.

Turbine vanes and blades are generally manufactured as single or double airfoil sections that must each be installed onto a turbine disk. Between each section, a gap at the endwalls through the blade passage is present, through which high pressure coolant is leaked. Furthermore, sections can become misaligned due to thermal expansion or centrifugal forces. Flow and heat transfer around the gap is complicated due to the interaction of the mainstream and the leakage flow. An experimental and computational study was undertaken to determine the physics of the leakage flow interaction for a realistic turbine blade endwall, and assess whether steady RANS CFD, commonly used for non-axisymmetric endwall design, can be used to accurately model this interaction. Computational models were compared against experimental observations of endwall heat transfer on a contoured endwall with a midpassage gap. Endwall heat transfer coefficients were determined experimentally by using infrared thermography to capture spatially-resolved surface temperatures on a uniform heat flux surface (heater) attached to the endwall. Predictions and measurements both indicated an increase in endwall heat transfer with increasing gap leakage flow, although the distribution of heat transfer coefficients along the gap was not well captured by CFD. A misalignment of the blade endwall causing a forward-facing step for the near-endwall flow resulted in a large highly turbulent recirculation region downstream of the step and high local heat transfer that was overpredicted by CFD. Conversely, a backward-facing step reduced turbulence and local heat transfer. The misprediction of local heat transfer around the gap is thought to be caused by unsteady interaction of the passage secondary flow and gap leakage flow, which cannot be well-captured by a steady RANS approach.

The accurate design, control and monitoring of the running gaps between static and moving components is vital to preserve the mechanical integrity and ensure the correct functioning of any compact rotating machinery. Throughout engine service, the rotor tip clearance undergoes large variations due to installation tolerances or as the result of different thermal expansion rates of the blades, rotor disk and casing during speed transients. Hence, active tip clearance control concepts and engine health monitoring systems rely on precise real-time gap measurements. Moreover, this tip gap information is crucial for engine development programs to verify the mechanical and aerothermal design, and validate numerical predictions. This paper presents an overview of the critical design requirements for testing engine-representative blade tip flows in a rotating turbine facility. The manuscript specifically focuses on the challenges related with the design, verification and monitoring of the running tip clearance during a turbine experiment. In the large-scale turbine facility of the von Karman Institute, a rainbow rotor was mounted for simultaneous aerothermal testing of multiple blade tip geometries. The tip shapes are a selection of high-performance squealer-like and contoured blade tip designs. On the rotor disc, the blades are arranged in seven sectors operating at different clearance levels from 0.5 up to 1.5% of the blade span. Prior to manufacturing, the blade geometry was modified to compensate for the radial deformation of the rotating assembly under centrifugal loads. A numerical procedure was implemented to minimize the residual unbalance of the rotor in rainbow configuration, and to optimize the placement of every single airfoil within each sector. Subsequently, the rotor was balanced in-situ to reduce the vibrations and satisfy the international standards for high balance quality. The single-blade tip clearance in rotation was measured by three fast-response capacitive probes located at three distinct circumferential locations around the rotor annulus. Additionally, the minimum running blade clearance is captured with wear gauges located at five axial positions along the blades chord. The capacitance probes are self-calibrated using a multi-test strategy at several rotational speeds. The in-situ calibration methodology and dedicated data reduction techniques allow the accurate measurement of the distance between the turbine casing and the local blade tip features (rims and cavities) for each rotating airfoil separately. General guidelines are given for the design and calibration of a tip clearance measurement system that meets the required measurement accuracy and resolution in function of the sensor uncertainty, nominal tip clearance levels and tip seal geometry.

In order to avoid the failure of bolted joints used on ultra-super critical steam turbine inner casing, their mechanical behavior should be strictly assessed, especially under the in-service conditions with steam temperature and pressure fluctuations. In this study, a 3D bolted inner casing is modeled using the commercial finite element software ABAQUS. The Norton-Bailey constitutive model and strain-life–based Manson-Coffin equation are applied to study the mechanical behavior of the bolted inner casing under in-service conditions. The results demonstrate that the inner casing could exert significant influence on the bolts by conduction, structure restriction and thermal expansion difference of the materials. The temperature and von Mises stress of the bolts experience continuous fluctuations during the steady state operation phase due to the constant variation in the steam temperature and pressure, and the alternative stresses give rise to fatigue damage of the material. The steam temperature and pressure fluctuations decline the fatigue lifetime of these high-temperature bolts in long-term operation, which should be taken into account in the design.

A numerical investigation of reacting flows in an advanced high-g cavity (HGC), Ultra-Compact Combustor (UCC) concept is conducted. The high-g cavity UCC (UCC-HGC) design uses high swirl in a circumferential cavity (CC) wrapped around a main stream annular flow. The high swirl is generated through angled CC driver jets. This centrifugal force is varied by changing the CC-to-core air mass flow ratio ( ṁ cc /ṁ core ) and jet inclination angle ( α jet ) relative to the cavity ring surface, while maintaining the global equivalence ratio (ϕ Global ) constant. Steady, rotational periodic, 3D simulations are performed following a multiphase, Reynolds-averaged Navier-Stokes (RANS), and non-premixed flamelet/progress variable (FPV) approach using a customized FLUENT. Results indicate that under non-reacting flow conditions the driver jets impose a very strong bulk swirl flow within the CC and the mainstream flow does not entrain into the CC. Thus, the maximum g-load is primarily sensitive to ṁ cc /ṁ core and secondarily to α jet . However, the g-loads become increasingly more sensitive to the latter at greater ṁ cc /ṁ core . Now, under reacting flow conditions, the flame interacts with the flow and the bulk swirl flow is diminished at low ṁ cc /ṁ core , while boosted at high ṁ cc /ṁ core . The former happens because the flame deflects the incoming driver jet flow, enhancing radial and axial velocity components (through thermal expansion), while diminishing the tangential flow velocity. This, in turn, weakens the g-loads within the CC to below its design g-load operation. On the other hand, at high ṁ cc /ṁ core and small α jet the flame is perpendicular to the bulk swirl flow, accelerating the flow tangential velocity and enhancing g-loads above its design operation. Qualitatively, the more and hotter the flame that can be sustained within the CC the shorter the flame length. The converse is also true. Flame length does not appear to be strongly influenced by ṁ cc /ṁ core and α jet . Even though g-loads appear to enhance reaction progress variable source (S C ) and, consequently, turbulent flame speed, through turbulence this does not necessarily mean that the turbulent flame speed under g-loads is various factors greater than its corresponding turbulent flame speed under 0g’s. As the ṁ cc /ṁ core increases the center-peaked radial temperature profile at intermediate α jet starts to deteriorate, whereas the radial temperature profile at low α jet improves. For high α jet , increasing ṁ cc /ṁ core has no substantial effect on the exit radial temperature profiles.

With increasing use of renewable energy sources, an industrial gas turbine is often a competitive solution to balance the power grid. However, life robustness approaches for gas turbine components operating under increasingly cyclic conditions, is a challenging task. Ductile superalloys, as Haynes 230, are often used in stationary gas turbine hot parts such as combustors. The main load for such components is due to non-homogeneous thermal expansion within or between parts. As the material is ductile there is considerable redistribution of stresses and strains due to inelastic deformations during the crack initiation phase. Therefore, the subsequent crack growth occurs through a material with significant residual stresses and strains. In this work, fatigue crack propagation experiments, including the initiation phase, have been performed on a single edge notched specimen under strain controlled conditions. The test results are compared to fracture mechanics analyses using the linear Δ K and the nonlinear Δ J approaches, and an attempt to quantify the difference in terms of a life prediction is made. For the tested notched geometry, material and strain ranges, the difference in the results using Δ K eff or Δ J eff are larger than the scatter seen when fitting the model to the experimental data. The largest differences can be found for short crack lengths, when the cyclic plastic work is the largest. The Δ J approach clearly shows better agreement with the experimental results in this regime.

Tighten force has much influence on tie-bolt fastened rotor dynamics. Temperature distribution in tie-bolt fastened rotor results in thermal expansion of rotor and rods. The difference of thermal expansion between rotor and rods causes the variation of bolt load. With considering the thermal contact conductance, the thermal model of tie-bolt fastened rotor was established by finite element method and the axial temperature distribution was obtained. The influences of surface roughness, nominal contact pressure and axial position of contact on axial temperature distribution were analysed. Based on temperature distribution in the tie-bolt fastened rotor, the variation of tighten force was investigated. Results show that nominal contact pressure, surface roughness and axial contact arrange have different influences on the variation of tighten force with temperature.