The inhomogeneity of temperature in a turbine is related to the nonuniform heat release and air injections in combustors, which is commonly measured by thermal couples at the turbine exit. Investigation of temperature inhomogeneity transportation in a multistage gas turbine should help in detecting and quantifying the over-temperature or flameout of combustors using turbine exhaust temperature. Here, the transportation of combustion inhomogeneity inside the four-stage turbine of a 300-MW gas turbine engine was numerically investigated. The computational domain included four turbine stages, consisting of more than 500 blades and vanes. Realistic components (N 2 , O 2 , CO 2 , and H 2 O) with variable heat capacities were considered for hot gas and cooling air. Coolants were added to the computational domain through more than 19,000 mass and momentum source terms. An unsteady Reynolds averaged Navier–Stokes computational fluid dynamics (URANS CFD) run with over-temperature/flameout at 6 selected combustors out of 24 was carried out. The temperature distributions at rotor–stator interfaces and the turbine outlet were quantified and characterized using Fourier transformations in the space domain. It is found that the transport process from the hot streaks/cold streaks at the inlet to the outlet is relatively stable. The cold and hot fluid is redistributed in time and space due to the stator and rotor blades, and in the region with a large parameter gradient at the inlet, a strong unsteady temperature field and a composition field appear. The distribution of the exhaust-gas composition has a stronger correlation with the inlet temperature distribution and is less susceptible to interactions.

The effect of the tip-shroud seal on the flutter onset of a shrouded turbine rotor blade, representative of a modern gas turbine, is numerically tested, and the contributions to the work per cycle of the aerofoil and the tip shroud are clearly identified. The numerical simulations are conducted using a linearized frequency-domain solver. The flutter stability of the shrouded rotor blade is evaluated for an edgewise mode and compared with the standard industrial approach of not including the tip-shroud cavity. It turns out that including the tip shroud significantly changes the stability prediction of the rotor blade. This is due to two facts. First, the amplitude of the unsteady pressure created in the inter-fin cavity due to the motion of the airfoil is much greater than that of the airfoil. The impact of this contribution increases with the frequency. Second, the effect of the outer shroud of the rotor blade, which usually is not included either in the simulations, has an opposite trend with the nodal diameter than the airfoil reducing the maximum and minimum damping. It is concluded that the combined effect of the seal and its platform tends to stabilize the edgewise mode of the rotor blade for all the examined nodal diameters and reduced frequencies. Finally, the numerical results are shown to be consistent with those obtained using an analytical simplified model to account for the effect of the labyrinth seals.

Jet engines with boundary layer ingestion (BLI) could offer significant reductions in aircraft fuel burn compared with podded turbofans. However, the engine fans must run continuously with severe inlet distortion, which is known to reduce stability. In this paper, an experimental study has been completed on a low-speed rig fan operating with a BLI-type inlet distortion. Unsteady casing static pressure measurements have been made at multiple locations during stall events. Steady-state, full-annulus area traverses have also been performed at rotor inlet and exit at a near-stall operating point. The reduction in stability caused by BLI is found to be small. It is found that with BLI the fan can operate stably despite the presence of localized regions where the rotor operating point lies beyond the stability boundary measured in clean flow. With the BLI-type distortion applied, the measured rotor incidence varies around the annulus due to nonuniform upstream velocity and swirl. The measured amplitude of unsteady casing pressure fluctuations just prior to stall is found to correlate with the circumferential variation of rotor incidence, suggesting that rotor incidence is a key variable affecting the creation and growth of flow disturbances. In regions of high incidence, disturbances resembling local flow separations are initiated. However, in regions of low or negative incidence, any disturbances decay rapidly. Full rotating stall with BLI occurs when high incidence regions are widespread enough to sustain disturbances which can propagate around the entire annulus.

The emerging renewable energy market calls for more advanced prediction tools for turbine transient operations in fast startup/shutdown cycles. Reliable numerical analysis of such transient cycles is complicated by the disparity in time scales of the thermal responses in fluid and solid domains. Obtaining fully coupled time-accurate unsteady conjugate heat transfer (CHT) results under these conditions would require to march in both domains using the time-step dictated by the fluid domain: typically, several orders of magnitude smaller than the one required by the solid. This requirement has strong impact on the computational cost of the simulation as well as being potentially detrimental to the accuracy of the solution due to accumulation of round-off errors in the solid. A novel loosely coupled CHT methodology has been recently proposed, and successfully applied to both natural and forced convection cases that remove these requirements through a source-term based modeling (STM) approach of the physical time derivative terms in the relevant equations. The method has been shown to be numerically stable for very large time steps with adequate accuracy. The present effort is aimed at further exploiting the potential of the methodology through a new adaptive time stepping approach. The proposed method allows for automatic time-step adjustment based on estimating the magnitude of the truncation error of the time discretization. The developed automatic time stepping strategy is applied to natural convection cases under long (2000 s) transients: relevant to the prediction of turbine thermal loads during fast startups/shutdowns. The results of the method are compared with fully coupled unsteady simulations showing comparable accuracy with a significant reduction of the computational costs.

A high speed 1–1/2 axial compressor stage is simulated in this paper using an unsteady Reynolds-averaged Navier–Stokes (URANS) solver for a full-annulus configuration to capture its nonsynchronous vibration (NSV) flow excitation with rigid blades. A third-order weighted essentially nonoscillatory scheme for the inviscid flux and a second-order central differencing for the viscous terms are used to resolve nonlinear unsteady fluid flows. A fully conservative rotor/stator sliding boundary condition (BC) is employed with multiple-processor capability for rotor/stator sliding interface that accurately captures unsteady wake propagation between the rotor and stator blades while conserving fluxes across the rotor/stator interfaces. The predicted dominant frequencies using the blade tip response signals are not harmonic to the engine order, which is the NSV excitation. The simulation is based on a rotor blade with a 1.1% tip-chord clearance. Comparison with the previous 1/7 annulus simulations show that the time-shifted phase-lag BCs used in the 1/7 annulus are accurate. For most of the blades, the NSV excitation frequency is 6.2% lower than the measurement in the rig test, although some blades displayed slightly different NSV excitation frequencies. The simulation confirms that the NSV is a full annulus phenomenon. The instability of the circumferential traveling vortices in the vicinity of the rotor tip due to the strong interaction of incoming flow is the main cause of the NSV excitation. This instability is present in all blades of the rotor annulus. For circumferentially averaged parameters like total pressure ratio, NSV is observed to have an effect on the radial profile, particularly at radial locations above 70% span. A design with a lower loading of the upper blade span and a higher loading of the midblade spans is recommended to mitigate or remove NSV.

In this paper, experiments and numerical modeling are used to quantify the effects of clearance and eccentricity on compressor performance and to examine the influence of each on flow distribution and stall margin. A change in the size of the tip-clearance gap influences the pressure rise and the stall margin of a compressor. Eccentricity of the tip-clearance gap then further exacerbates the negative effects of increasing tip-clearance. There are few studies in the literature dealing with the combined effect of clearance and eccentricity. There is also little guidance for engine designers, who have traditionally used rules of thumb to quantify these effects. One such rule states that the stall margin of an eccentric machine will be equal to that of a concentric machine with uniform clearance equal to the maximum eccentric clearance. In this paper, this rule of thumb is checked using experimental data and found to be overly pessimistic. In addition, eccentric clearance causes a variation in axial velocity around the circumference of the compressor. The current study uses a three-dimensional model which demonstrates the importance of radial flow gradients in capturing this redistribution. Flow redistribution has been treated analytically in the past, and for this reason, previous modeling has been restricted to two dimensions. The circumferential variation in axial velocity is also examined in terms of the local stability of the flow by considering the stalling flow coefficient of an equivalent axisymmetric compressor with the same local tip-clearance. The large clearance sector of the annulus is found to operate beyond its equivalent axisymmetric stall limit, which means that the small clearance sector of the annulus must be stabilizing the large clearance sector. An improved rule of thumb dealing with the effects of eccentricity is presented.

In the present paper, the results of an experimental and numerical investigation of the hub cavity modes and their migration into the main annulus flow are presented. A one-and-a-half stage, unshrouded and highly loaded axial turbine configuration with three-dimensionally shaped blades and cylindrical end walls has been tested in an axial turbine facility. Both the blade design and the rim seal purge flow path are representative to modern high-pressure gas turbines. The unsteady flow field at the hub cavity exit region has been measured with the fast-response aerodynamic probe (FRAP) for two different rim seal purge flow rates. Furthermore, fast-response wall-mounted pressure transducers have been installed inside the cavity. Unsteady full-annular computational fluid dynamics (CFD) simulations have been employed in order to complement the experimental work. The time-resolved pressure measurements inside the hub cavity reveal clear cavity modes, which show a strong dependency on the injected amount of rim seal purge flow. The numerical predictions provide information on the origin of these modes and relate them to pronounced ingestion spots around the circumference. The unsteady probe measurements at the rim seal interface show that the signature of the hub cavity induced modes migrates into the main annulus flow up to 30% blade span. Based on that, an aerodynamic loss mechanism has been found, showing that the benefit in loss reduction by decreasing the rim seal purge flow rate is weakened by the presence of turbine hub cavity modes.

The coupling between the bleed system and the flowfield of a downstream compressor stage is studied using two approaches. In the first approach, three-dimensional, full annulus, unsteady computations simulate the flow in a low-speed research compressor with nonuniform bleed extraction. Comparisons with experimental data show that the flow prediction in the main annulus is accurate to within 0.005 of flow coefficient and 0.5 deg of flow angle. The computational fluid dynamics (CFD) is then used to provide a description of flow within the bleed system itself. In the second approach, a two-dimensional mean radius model, similar to that adopted by Hynes and Greitzer in the previous work on compressor stability, is used to simulate the response of the compressor to nonuniform bleed. This model is validated against experimental data for a single-stage compressor, and despite the inherent assumptions (two-dimensional flow and simplified compressor response), provides a satisfactory prediction of the flow for preliminary design purposes with orders of magnitude less computational cost than full 3D CFD. The model is then used to investigate the effect of different levels of bleed nonuniformity and of varying the axial distance between the bleed and the downstream stage. Reducing bleed nonuniformity and moving the stage away from the bleed slot are predicted to reduce the circumferential nonuniformity of the flow entering the stage.

In the present study, an application for efficient cooling of turbine liner segments employing pulsating impinging jets was investigated. A combined numerical and experimental study was conducted to evaluate the design of a case cavity device which utilizes the periodically unsteady pressure distribution caused by the rotor blades to excite a pulsating impinging jet. Through an opening between the main annulus and a case cavity, pressure pulses from the rotor blades propagated into this cavity and caused a strong pressure oscillation inside. The unsteady computational fluid dynamics (CFD) results were in good qualitative agreement with the measurement data obtained using high-frequency pressure transducers and hot wire anemometry. Furthermore, the numerical study revealed the formation of distinct toroidal vortex structures at the nozzle outlet as a result of the jet pulsation. Within the scope of the measurements, the influence of the operating point on the pressure propagation inside the cavity was investigated. The dependence of shape and amplitude of the pressure oscillation on engine speed and stage pressure ratio was found to be in accordance with an analytical consideration.

This paper describes experiments in a subscale axial turbine stage equipped with an axially overlapping radial-clearance seal at the disk cavity rim and a labyrinth seal radially inboard which divides the disk cavity into a rim cavity and an inner cavity. An orifice model of the rim seal is presented; values of ingestion and egress discharge coefficients based on the model and experimental data are reported for a range of cavity purge flow rate. In the experiments, time-averaged pressure distribution was measured in the main gas annulus and in the disk cavity; also measured was the time-averaged ingestion into the cavity. The pressure and ingestion data were combined to obtain the discharge coefficients. Locations on the vane platform 1 mm upstream of its lip over two vane pitches circumferentially defined the main gas annulus pressure; in the rim cavity, locations at the stator surface in the radially inner part of the “seal region” over one vane pitch defined the cavity pressure. For the sealing effectiveness, two locations in the rim cavity at the stator surface, one in the “mixing region” and the other radially further inward at the beginning of the stator boundary layer were considered. Two corresponding sets of ingestion and egress discharge coefficients are reported. The ingestion discharge coefficient was found to decrease in magnitude as the purge flow rate increased; the egress discharge coefficient increased with purge flow rate. The discharge coefficients embody fluid-mechanical effects in the ingestion and egress flows. Additionally, the minimum purge flow rate required to prevent ingestion was estimated for each experiment set and is reported. It is suggested that the experiments were in the combined ingestion (CI) region with externally induced (EI) ingestion being the dominant contributor.

In gas turbines, rim seals are fitted at the periphery of the wheel-space between the turbine disk and its adjacent casing; their purpose is to reduce the ingress of hot mainstream gases. A superposed sealant flow, bled from the compressor, is used to purge the wheel-space or at least dilute the ingress to an acceptable level. The ingress is caused by the circumferential variation of pressure in the turbine annulus radially outward of the seal. Engine designers often use double-rim seals where the variation in pressure is attenuated in the outer wheel-space between the two seals. This paper describes experimental results from a research facility that models an axial turbine stage with engine-representative rim seals. The radial variation of CO2 gas concentration, swirl, and pressure, in both the inner and outer wheel-space, are presented over a range of purge flow rates. The data are used to assess the performance of two seals: a datum double-rim seal and a derivative with a series of radial fins. The concept behind the finned seal is that the radial fins increase the swirl in the outer wheel-space; measurements of swirl show the captive fluid between the fins rotate with near solid body rotation. The improved attenuation of the pressure asymmetry, which governs the ingress, results in an improved performance of the inner geometry of the seal. The fins also increased the pressure in the outer wheel-space and reduced the ingress though the outer geometry of the seal.

This paper presents an industrial perspective on the potential use of multiple-airfoil row unsteady computational fluid dynamics (CFD) calculations in high-pressure turbine design cycles. A sliding-mesh unsteady CFD simulation is performed for a high-pressure turbine section of a modern aviation engine at conditions representative of engine take-off. The turbine consists of two stages plus a center-frame strut upstream of the low-pressure turbine. The airfoil counts per row are such that a half-annulus model domain must be simulated for periodicity. The total model domain size is 170 MM computational grid points and the solution requires approximately nine days of clock time on 6288 processing cores of a Cray XE6 supercomputer. Airfoil and endwall cooling flows are modeled via source term additions to the flow. The endwall flowpath cavities and their purge/leakage flows are resolved in the computational meshes to an extent. The time-averaged temperature profile solution is compared with static rake data taken in engine tests. The unsteady solution shows a considerable improvement in agreement with the rake data, compared with a steady-state solution using circumferential mixing planes. Passage-to-passage variations in the gas temperature prediction are present in the 2nd stage, due to nonperiodic alignment between the nozzle vanes and rotor blades. These passage-to-passage differences are quantified and contrasted.

This paper summarizes the work of a five year research program into the heat transfer within cavities adjacent to the main annulus of a gas turbine. The work has been a collaboration between several gas turbine manufacturers, also involving a number of universities working together. The principal objective of the study has been to develop and validate computer modeling methods of the cooling flow distribution and heat transfer management, in the environs of multistage turbine disk rims and blade fixings, with a view to maintaining component and subsystem integrity, while achieving optimum engine performance and minimizing emissions. A fully coupled analysis capability has been developed using combinations of commercially available and in-house computational fluid dynamics (CFD) and finite element (FE) thermomechanical modeling codes. The main objective of the methodology is to help decide on optimum cooling configurations for disk temperature, stress, and life considerations. The new capability also gives us an effective means of validating the method by direct use of disk temperature measurements, where otherwise, additional and difficult to obtain parameters, such as reliable heat flux measurements, would be considered necessary for validation of the use of CFD for convective heat transfer. A two-stage turbine test rig has been developed and improved to provide good quality thermal boundary condition data with which to validate the analysis methods. A cooling flow optimization study has also been performed to support a redesign of the turbine stator well cavity to maximize the effectiveness of cooling air supplied to the disk rim region. The benefits of this design change have also been demonstrated on the rig. A brief description of the test rig facility will be provided together with some insights into the successful completion of the test program. Comparisons will be provided of disk rim cooling performance for a range of cooling flows and geometry configurations. The new elements of this work are the presentation of additional test data and validation of the automatically coupled analysis method applied to a partially cooled stator well cavity (i.e., including some local gas ingestion) and also the extension of the cavity cooling design optimization study to other new geometries.

In core gas turbines relatively cold air is purged through the hub gap between the stator and rotor in order to seal the disk space against flow ingestion from the main annulus. Although the sealing mass flow rate is commonly very small compared to the main annulus mass flow rate, it can have significant effects on the development of the passage endwall flows and on the overall loss generation. In this paper, the interaction between the annulus and rim sealing flows is investigated using numerical simulations of a generic high-pressure turbine. At first, the numerical approach is validated by comparing the results of calculations to measurement data at the design flow conditions. Following that, results from steady and unsteady calculations are used to describe in detail the aerodynamics in overlap-type rim seals and their effects on the blade passage flow. It is found that the flow interaction at the rim seal interface is strongly influenced by the velocity deficit of the rim sealing flow relative to the annulus flow as well as by the circumferentially nonuniform pressure field imposed by the rotor blades. At typical sealing flow conditions, the flow interaction at the rim seal interface is found to be naturally unsteady, with periodical vortex shedding into the rotor passage. The rim seal geometry, in particular close to the annulus interface, is found to have an effect on the frequency of the shedding process. Finally, it is shown that the rim seal flow interaction can be stabilized by increasing either the sealing mass flow rate or the sealing tangential velocity.

This paper describes experimental results from a research facility, which experimentally models hot gas ingress into the wheel-space of an axial turbine stage. Measurements of CO2 gas concentration in the rim-seal region and inside the wheel-space are used to assess the performance of generic (though engine-representative) single and double seals in terms of the variation of concentration effectiveness with sealing flow rate. The variation of pressure in the turbine annulus, which governs externally induced ingress, was obtained from steady pressure measurements downstream of the vanes. The benefit of using double seals is demonstrated: the ingested gas is shown to be predominately confined to the outer wheel-space radially outward of the inner seal; and in the inner wheel-space, radially inward of the inner seal, the effectiveness is shown to be significantly higher. Criteria for ranking the performance of single and double seals are proposed, and the performance limit for any double seal is shown to be one in which the inner seal is exposed to rotationally induced ingress. Although the ingress is a consequence of an unsteady, three-dimensional flow field and the cause-effect relationship between pressure and the sealing effectiveness is complex, the experimental data is shown to be successfully calculated by simple effectiveness equations developed from a theoretical model. The data illustrate that, for similar turbine-stage velocity triangles, the effectiveness can be correlated using two empirical parameters. In principle, these correlations could be extrapolated to a geometrically similar turbine operating at engine-representative conditions.

The spoiling effects of rim seal flow are studied at the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt. Detailed flow field measurements and efficiency measurements were performed for various ingress and egress setups and will be presented in this paper. Efficiency measurements show an efficiency decrease as the rim seal mass flow is increased. Five hole probe measurements upstream and downstream of the second stator row show that an increasing rim seal mass flow leads to an increased pressure loss across the stator, to altered incidence angles and to an intensification of secondary flow structures within the lower 50% span. Static pressure taps at the stator profile primarily show altered aerodynamic loading with increased rim seal air. In addition, the end wall profile pressure was measured at the stator 2 hub. It can be seen that seal air injection causes increased pressure fluctuations on the platform. Temperature measurements with a temperature difference between rim seal and main annulus flow show that rim seal air primarily enters the passage vortex.

Reliable means of predicting heat transfer in cavities adjacent to the main gas path are increasingly being sought by engineers involved in the design of gas turbines. In this paper, an interim summary of the results of a five-year research program sponsored by the European Union (EU) and several leading gas turbine manufacturers and universities will be presented. Extensive use is made of computational fluid dynamics (CFD) and finite element (FE) modeling techniques to understand the thermo-mechanical behavior of a turbine stator well cavity, including the interaction of cooling air supply with the main annulus gas. The objective of the study has been to provide a means of optimizing the design of such cavities for maintaining a safe environment for critical parts, such as disc rims and blade fixings, while maximizing the turbine efficiency and minimizing the fuel burn and emissions penalties associated with the secondary airflow system. The modeling methods employed have been validated against data gathered from a dedicated two-stage turbine rig running at engine representative conditions. Extensive measurements are available for a range of flow conditions and alternative cooling arrangements. The analysis method has been used to inform a design change, which is also to be tested. Comparisons are provided between the predictions and measurements of the turbine stator well component temperature.

This paper describes a new research facility which experimentally models hot gas ingestion into the wheel-space of an axial turbine stage. Measurements of the CO2 gas concentration in the rim-seal region and inside the cavity are used to assess the performance of two generic (though engine-representative) rim-seal geometries in terms of the variation of concentration effectiveness with sealing flow rate. The variation of pressure in the turbine annulus, which governs this externally-induced (EI) ingestion, was obtained from steady pressure measurements downstream of the vanes and near the rim seal upstream of the rotating blades. Although the ingestion through the rim seal is a consequence of an unsteady, three-dimensional flow field and the cause-effect relationship between the pressure and the sealing effectiveness is complex, the experimental data is shown to be successfully calculated by simple effectiveness equations developed from a previously published orifice model. The data illustrate that, for similar turbine-stage velocity triangles, the effectiveness can be correlated using a nondimensional sealing parameter, Φo. In principle, and within the limits of dimensional similitude, these correlations should apply to a geometrically-similar engine at the same operating conditions. Part II of this paper describes an experimental investigation of rotationally-induced (RI) ingress, where there is no mainstream flow and consequently no circumferential variation of external pressure.

In recent papers, orifice models have been developed to calculate the amount of ingestion, or ingress, that occurs through gas-turbine rim seals. These theoretical models can be used for externally induced (EI) ingress, where the pressure differences in the main gas path are dominant, and for rotationally induced (RI) ingress, where the effects of rotation in the wheel space are dominant. Explicit “effectiveness equations,” derived from the orifice models, are used to express the flow rate of sealing air in terms of the sealing effectiveness. These equations contain two unknown terms: Φmin, a sealing flow parameter, and Γc, the ratio of the discharge coefficients for ingress and egress. The two unknowns can be determined from concentration measurements in experimental rigs. In this paper, maximum likelihood estimation is used to fit the effectiveness equations to experimental data and to determine the optimum values of Φmin and Γc. The statistical model is validated numerically using noisy data generated from the effectiveness equations, and the simulated tests show the dangers of drawing conclusions from sparse data points. Using the statistical model, good agreement between the theoretical curves and several sets of previously published effectiveness data is achieved for both EI and RI ingress. The statistical and theoretical models have also been used to analyze previously unpublished experimental data, the results of which are included in separate papers. It is the ultimate aim of this research to apply the effectiveness data obtained at rig conditions to engine-operating conditions.

Gas turbine engines operating in a hostile environment, polluted with sand or dust particles, are susceptible to erosion damage, mostly at the front axial fans and compressors. Accurately predicting the erosion pattern and rate due to sand ingestion is one of the major challenges faced by the transportation and power industries. Maintenance costs are scrutinized and intensive research efforts are currently deployed in predictive life assessment tools to minimize the overhaul down time. The conventional prediction methods were usually based on steady-state simulations of gas-phase flows through a single blade passage per blade row to reduce the computational cost. However, the multistage turbomachinery flows are intrinsically subject to unsteadiness, especially due to stator-rotor interactions, which may affect sand particle trajectories even if a one-way coupling method is considered. Furthermore, an unsteady stator-rotor interaction requires a whole-annulus model at great computational cost to avoid simplifications of the geometries or flow physics. To study the effects of the stator-rotor interaction on sand particle trajectories and erosion, an axial fan with inlet guide vanes is investigated, based on the whole annulus computations of both steady and unsteady gas-phase flows, each of which is then followed by a Lagrangian particle tracking step. A numerical algorithm for tracking particles driven by the unsteady gas-phase flow is presented. The comparison of the numerical predictions with the experimental data confirms the validity and necessity of the unsteady computational fluid dynamics (CFD) model in providing adequate predictions of sand erosion in the axial fan.