Showing papers on "Tip clearance published in 2004"
TL;DR: In this article, the authors describe the nature of 3D separation and address the way in which topological rules based on a linear treatment of the Navier-Stokes equations can predict properties of the limiting streamlines, including the singularities which form.
Abstract: Flow separations in the corner regions of blade passages are common. The separations are three dimensional and have quite different properties from the two-dimensional separations that are considered in elementary courses of fluid mechanics. In particular the consequences for the flow may be less severe than the two-dimensional separation. This paper describes the nature of three-dimensional separation and addresses the way in which topological rules, based on a linear treatment of the Navier-Stokes equations, can predict properties of the limiting streamlines, including the singularities which form. The paper shows measurements of the flow field in a linear cascade of compressor blades and compares these with the results of 3D CFD. For corners without tip clearance, the presence of three-dimensional separation appears to be universal and the challenge for the designer is to limit the loss and blockage produced. The CFD appears capable of predicting this.Copyright © 2004 by ASME
225 citations
TL;DR: In this article, the noise due to tip clearance (TC) flow in axial flow fans operating at a design and off-design conditions is analyzed by an experimental measurement using two hot-wire probes rotating with the fan blades.
109 citations
01 Jan 2004
TL;DR: In this article, the role of tip clearance flow and its interaction with the passage shock on stall inception is analyzed in detail, and the numerically obtained flow fields are interrogated to identify the roles of flow interactions between the tip-clearing flow, the passage-surface boundary layer and the blade/endwall boundary layers.
Abstract: The current paper reports on investigations aimed at advancing the understanding of the flow field near the casing of a forward-swept transonic compressor rotor. The role of tip clearance flow and its interaction with the passage shock on stall inception are analyzed in detail. Steady and unsteady three-dimensional viscous flow calculations are applied to obtain flow fields at various operating conditions. The numerical results are first compared with available measured data. Then, the numerically obtained flow fields are interrogated to identify the roles of flow interactions between the tip clearance flow, the passage shock, and the blade/endwall boundary layers. In addition to the flow field with nominal tip clearance, two more flow fields are analyzed in order to identify the mechanisms of blockage generation: one with zero tip clearance, and one with nominal tip clearance on the forward portion of the blade and zero clearance on the aft portion. The current study shows that the tip clearance vortex does not break down, even when the rotor operates in a stalled condition. Interaction between the shock and the suction surface boundary layer causes the shock, and therefore the tip clearance vortex, to oscillate. However, for the currently investigated transonic compressor rotor, so-called breakdown of the tip clearance vortex does not occur during stall inception. The tip clearance vortex originates near the leading edge tip, but moves downward in the spanwise direction inside the blade passage. A low momentum region develops above the tip clearance vortex from flow originating from the casing boundary layer. The low momentum area builds up immediately downstream of the passage shock and above the core vortex. This area migrates toward the pressure side of the blade passage as the flow rate is decreased. The low momentum area prevents incoming flow from passing through the pressure side of the passage and initiates stall inception. It is well known that inviscid effects dominate tip clearance flow. However, complex viscous flow structures develop inside the casing boundary layer at operating conditions near stall.Copyright © 2004 by ASME
83 citations
TL;DR: In this article, the authors presented a numerical parametric study of tip clearance coupled with casing treatment for a transonic axial-flow compressor NASA Rotor 37 and found that the casing treatments were an effective means of reducing the negative effects of tip gap flow and vortex, resulting in improved performance and stability.
Abstract: The control of tip leakage flow (TLF) through the clearance gap between the moving and stationary components of rotating machines is still a high-leverage area for improvement of stability and performance of aircraft engines. Losses in the form of flow separation, stall, and reduced rotor work efficiency are results of the tip leakage vortex (TLV) generated by interaction of the main flow and the tip leakage jet induced by the blade pressure difference. The effects are more detrimental in transonic compressors due to the interaction of shock-TLV. It has been previously shown that the use of slots and grooves in the casing over tip of the compressor blades, known as casing treatment, can substantially increase the stable flow range and therefore the safety of the system but generally with some efficiency penalties. This paper presents a numerical parametric study of tip clearance coupled with casing treatment for a transonic axial-flow compressor NASA Rotor 37. Compressor characteristics have been compared to the experimental results for smooth casing with a 0.356 mm tip clearance and show fairly good agreement. Casing treatments were found to be an effective means of reducing the negative effects of tip gap flow and vortex, resulting in improved performance and stability. The present work provides guidelines for improvement of steady-state performance of the transonic axial-flow compressors and improvement of the stable operating range of the system.Copyright © 2004 by ASME
79 citations
TL;DR: In this paper, the effect of tip clearance on the cooling performance of the microchannel heat sink is presented under the fixed pumping power condition, and the thermal resistance is defined for evaluating its cooling performance.
72 citations
TL;DR: In this paper, the effect of the rim height and the tip gap clearance on the heat transfer coefficients on the blade tip and near tip regions were measured with two different rim geometries.
62 citations
TL;DR: In this article, a 3-D model with tip clearance has been generated to predict the internal flow and performance of the turbine and it is found that the comparison between computed and experimental data is good, quantitatively and qualitatively.
62 citations
01 Jan 2004
TL;DR: In this paper, the effect of tip clearance on the transient process of rotating stall evolution has been studied experimentally in a low-speed axial compressor stage with various stator-rotor gaps.
Abstract: Effect of the tip clearance on the transient process of rotating stall evolution has been studied experimentally in a low-speed axial compressor stage with various stator-rotor gaps. In the previous authors’ experiments for the small tip clearance, the stall evolution process of the rotor was sensitive to the gaps between the blade rows. For the large tip clearance, however, little difference is observed in the evolution processes independently of the blade row gap. In the first half process, it is characterized by gradual reduction of overall pressure-rise with flow rate decreasing, and the number of short length-scale disturbances is increasing with their amplitude increasing. In the latter half a long length-scale disturbance develops rapidly to result in deep stall. Just before the stall inception the spectral power density of the casing wall pressure reveals the existence of rotating disturbances with broadband high frequency near a quarter of the blade passing frequency. This is caused by the short length-scale disturbances occurring intermittently. A flow model is presented to explain mechanisms of the rotating short length-scale disturbance, which includes a tornado-like separation vortex and tip-leakage vortex breakdown. The model is supported by a result of a numerical unsteady flow simulation.Copyright © 2004 by ASME
55 citations
Patent•
21 Dec 2004
TL;DR: In this article, a system for controlling blade tip clearance in a turbine is presented, which includes a stator including a shroud having a plurality of shroud segments (38) and a rotor (12) including a blade rotatable within the shroud.
Abstract: A system for controlling blade tip clearance in a turbine. The system includes a stator including a shroud having a plurality of shroud segments (38) and a rotor (12) including a blade (14) rotatable within the shroud. An actuator assembly (30) is positioned radially around the shroud and includes a plurality of actuators (18). A sensor (16) senses a turbine parameter and generates a sensor signal representative of the turbine parameter. A modeling module (22) generates a tip clearance prediction in response to turbine cycle parameters. A controller (20) receives the sensor signal and the tip clearance prediction and generates at least one command signal. The actuators (18) include at least one actuator receiving the command signal and adjusts a position of at least one of the shroud segments (38) in response to the command signal.
42 citations
TL;DR: In this paper, the impact of forward swept rotors on axial compressors was investigated and two different configurations were examined, one with strong tip-clearance flows and the other with more moderate levels.
Abstract: This paper presents an experimental and analytical study of the impact of forward swept rotors on tip-limited, low-speed, multistage axial compressors. Two different configurations were examined, one with strong tip-clearance flows and the other with more moderate levels. Evaluations were done at multiple rotor tip clearances to assess differences in clearance sensitivity. Compared to conventionally stacked radial rotors, the forward swept blades demonstrated improvements in stall margin, efficiency and clearance sensitivity. The benefits were more pronounced for the configuration with stronger tip-clearance flows. Detailed flow measurements and three-dimensional viscous CFD analyses were used to investigate the responsible flow mechanisms. Forward sweep causes a spanwise redistribution of flow toward the blade tip and reduces the tip loading in terms of static pressure coefficient. This results in reduced tip-clearance flow blockage, a shallower (more axial) vortex trajectory and a smaller region of reversed flow in the clearance gap.
41 citations
Journal Article•
01 Jan 2004
TL;DR: In this article, the authors measured the natural and forced responses of the tip clearance vortex in a linear compressor cascade to characterize periodic unsteadiness of the vortex and suggested a physical explanation of the source of the observed periodic instability.
Abstract: Natural and forced responses of the tip clearance vortex are measured in a linear compressor cascade to characterize periodic unsteadiness of the tip clearance vortex. There exists a natural frequency at which the tip clearance vortex is the most receptive to external forcing, thus resulting in mixing enhancement and flow blockage reduction. A physical explanation of the source of the observed periodic unsteadiness is suggested based on the trailing vortex instability theory. Observations of the time scale for the unsteadiness from different compressor geometries and flow conditions are shown to scale with a reduced frequency based on convective time through the blade passage.Copyright © 2004 by ASME
01 Jan 2004
TL;DR: In this paper, the secondaire reference record was modified on 2016-08-08 to include the refroidisation of the second-order reference record of the first-order refroidizer.
Abstract: Keywords: jeu ; transfert de chaleur ; refroidissement ; ecoulement : secondaire Reference Record created on 2005-11-18, modified on 2016-08-08
01 Jul 2004
TL;DR: In this article, a control system for active turbine tip clearance control in a generic commercial turbofan engine through design and analysis is presented. The control objective is to articulate the shroud in the high pressure turbine section in order to maintain a certain clearance set point given several possible engine transient events.
Abstract: This paper addresses the requirements of a control system for active turbine tip clearance control in a generic commercial turbofan engine through design and analysis. The control objective is to articulate the shroud in the high pressure turbine section in order to maintain a certain clearance set point given several possible engine transient events. The system must also exhibit reasonable robustness to modeling uncertainties and reasonable noise rejection properties. Two actuators were chosen to fulfill such a requirement, both of which possess different levels of technological readiness: electrohydraulic servovalves and piezoelectric stacks. Identification of design constraints, desired actuator parameters, and actuator limitations are addressed in depth; all of which are intimately tied with the hardware and controller design process. Analytical demonstrations of the performance and robustness characteristics of the two axisymmetric LQG clearance control systems are presented. Takeoff simulation results show that both actuators are capable of maintaining the clearance within acceptable bounds and demonstrate robustness to parameter uncertainty. The present model-based control strategy was employed to demonstrate the tradeoff between performance, control effort, and robustness and to implement optimal state estimation in a noisy engine environment with intent to eliminate ad hoc methods for designing reliable control systems.
TL;DR: In this article, a compressor cascade test rig consisting of seven prismatic controlled diffusion blades with the middle blade being oscillated in a three-dimensional bending/flapping mode was developed.
Abstract: An experiment is carried out to enhance the understanding of three-dimensional blade aeroelastic mechanisms and to produce test data for validation of numerical methods. A compressor cascade test rig developed consists of seven prismatic controlled diffusion blades with the middle blade being oscillated in a three-dimensional bending/ flapping mode. Steady and unsteady pressure measurements are performed at six spanwise sections for three reduced frequencies and two tip-clearance gaps. Off-board pressure transducers are utilized with a transfer-function method correcting tubing errors. The tuned cascade results are constructed by using the influence coefficient method. The results illustrate fully three-dimensional unsteady behaviour. The blades are aeroelastically destabilized as the tip gap is increased, and the destabilizing effect of the tip-clearance influences most of the span. The total aerodynamic damping at the least stable interblade phase angle is reduced by 27% when the tip gap is increased from nearly 0 to 2% span.
10 Feb 2004
TL;DR: Detailed measurements in compressor cascades, tests in a low-speed single stage Deverson compressor rig and considerable use of Denton multi-stage Reynolds averaged Navier-Stokes (RANS) solver MULTIP have provided better understanding and improved predictive capability of this important phenomenon.
Abstract: A lack of understanding of the nature and characteristics of three-dimensional (3D) separation, which is inherent in compressor blade passages and tip clearance dominated 3D flow in the rotor tip region, has limited the success of 3D blade design. This dissertation is therefore aimed at contributing to compressor design process by exploring the formation of 3D separations. Detailed measurements in compressor cascades, tests in a low-speed single stage Deverson compressor rig and considerable use of Denton multi-stage Reynolds
averaged Navier-Stokes (RANS) solver MULTIP have provided better understanding and improved predictive capability of this important phenomenon. An attempt has also been made to clarify the role of different flow mechanisms.
Tests carried out to determine the effect of surface roughness in the single-stage axial compressor show that stage performance, even at design point is extremely sensitive to roughness around the leading edge and peak suction regions because of the effect on 3D separations. A numerical model to simulate the effect of roughness was formulated and incorporated into MULTIP and this showed good agreement with measurements.
A method of estimating the thickness of the 3D separated layer has been developed, which was used in a parametric study to assess the sensitivity of 3D separations to key flow and design parameters in axial compressors. Investigation into endwall flow control methods was undertaken, based on the understanding of the formation of 3D separation. These include clearance flows, endwall fences and endwall dividing streamline suction. The latter was demonstrated to be the most effective method of control.
The knowledge gained in predicting 3D separation was then used to explore the 3D nature of the flow around rotor tip sections and how this is influenced by key 3D design techniques. The performance of the existing 3D rotor and the modified version with tip chord extension was analysed using detailed CFD, with carefully refined mesh especially in the blade tip/clearance region. This approach enabled key problems with the oliginal 3D
design to be diagnosed. The lesson learned from the performance of the 3D rotor therefore guided a re-design of the rotor tip. It was found that a more modest amount of blade lean is enough to achieve improved performance improvement in the tip region.
Suggested future work that will aid further understanding of 3D separations in terms of
flow mechanism, unsteadiness and control are also presented.
TL;DR: A powerful computational technique, large-eddy simulation, has been used to study the detailed flow dynamics in the tip-gap region of hydraulic turbomachines as discussed by the authors, which can lead to reduced performance, increased noise, and structural vibration and erosion.
Abstract: A powerful computational technique, large-eddy simulation, helps researchers study the detailed flow dynamics in the tip-gap region of hydraulic turbomachines. LES also helps researchers investigate ways to mitigate undesirable effects, such as cavitation, which can lead to reduced performance, increased noise, and structural vibration and erosion.
TL;DR: In this paper, a unique comparative experimental and numerical investigation carried out on two test cases with shroud configurations differing only in the labyrinth seal path, is presented in order to analyze the flow effects reflected on the overall performance in a multi-stage environment.
Abstract: A unique comparative experimental and numerical investigation carried out on two test cases with shroud configurations differing only in the labyrinth seal path, is presented in this paper. The blade geometry and tip clearance is identical in the two test cases. The geometries under investigation are representative of an axial turbine with a full and partial shroud, respectively. Global performance and flow field data were acquired and analyzed. Computational simulations were carried out to complement the investigation and to facilitate the analysis of the steady and unsteady flow measurements. A detailed comparison between the two test cases is presented in terms of flow field analysis and performance evaluation. The analysis focuses on the flow effects reflected on the overall performance in a multi-stage environment. Strong interaction between the cavity flow and the blade tip region of the rotor blades is observed up to the blade mid span. A marked effect of this interaction can be seen in the downstream second stator where different vortex structures are observed. Moreover, in the partial shroud test case, a strong tip leakage vortex is developed from the first rotor and transported through the downstream blade row. A measurable change in the second stage efficiency was observed between the two test cases. In low aspect ratio blades within a multistage environment, small changes in the cavity geometry can have a significant effect on the mainstream flow. The present analysis has shown that an integrated and matched blade-shroud aerodynamic design has to be adopted to reach optimal performances. The additional losses resulting from small variations of the sealing geometry could result in a gain of up to one point in the overall stage efficiency.Copyright © 2004 by ASME
TL;DR: In this paper, the authors quantitatively estimate the tip clearance effect on the performance drop and the efficiency drop of a centrifugal compressor with six different tip clearances, and the additional entropy generation was modeled with all the kinetic energy of the tip leakage flow.
Abstract: Numerical simulations have been performed to investigate tip clearance effect on through-flow and performance of a centrifugal compressor which has the same configuration of impeller with six different tip clearances. Secondary flow and loss distribution have been surveyed to understand the flow mechanism due to the tip clearance. Tip leakage flow strongly interacts with mainstream flow and considerably changes the secondary flow and the loss distribution inside the impeller passage. A method has been described to quantitatively estimate the tip clearance effect on the performance drop and the efficiency drop. The tip clearance has caused specific work reduction and additional entropy generation. The former, which is called inviscid loss, is independent of any internal loss and the latter, which is called viscous loss, is dependent on every loss in the flow passage. Two components equally affected the performance drop as the tip clearances were small, while the efficiency drop was influenced by the viscous component alone. The additional entropy generation was modeled with all the kinetic energy of the tip leakage flow. Therefore, the present paper can provide how to quantitatively estimate the tip clearance effect on the performance and efficiency.
01 Jan 2004
TL;DR: In this paper, a numerical study has been performed to simulate the tip leakage flow and heat transfer on the first stage of a high-pressure turbine, which represents a modern gas turbine blade geometry.
Abstract: in Undetermined A numerical study has been performed to simulate the tip leakage flow and heat transfer on the first stage of a high-pressure turbine, which represents a modern gas turbine blade geometry. The low Re k-ω (SST) model is used to model the turbulence. Calculations are performed for both a flat and a squealer blade tip for three different tip gap clearances. The computations were carried out using a single blade with periodic conditions imposed along the boundaries in the circumferential (pitch) direction. The predicted tip heat transfer and static pressure distributions show reasonable agreement with experimental data. It was also observed that the tip clearance has a significant influence on local tip heat transfer coefficient distribution. The flat tip blade provides a higher overall heat transfer coefficient than the squealer tip blade. (Less)
Book•
01 Jan 2004
01 Jan 2004
TL;DR: In this article, a tip and seal clearance model for transient engine performance has been developed to be used in transient synthesis codes, where the model parameters have been identified from thermo-mechanical finite element models and the model calculates symmetric rotor tip clearance changes and symmetric seal clearance changes in the secondary air system for engine transients.
Abstract: Secondary effects, such as heat transfer from fluid to engine structure and the resulting changes in tip and seal clearances affect component performance and stability. A tip clearance model to be used in transient synthesis codes has been developed. The tip clearance model is derived as a state space structure. The model parameters have been identified from thermo-mechanical finite element models. The model calculates symmetric rotor tip clearance changes in the turbo-machinery and symmetric seal clearance changes in the secondary air system for engine transients within the entire flight envelope. The resulting changes in efficiency, capacity and cooling air flows are fed into the performance program. Corrections for tip clearance changes on the component characteristics are derived from rig tests. The effect of seal clearance changes on the secondary air system is derived using sophisticated air system models. The clearance model is validated against FE thermo-mechanical models. The modeling method of modifying the component characteristics is verified comparing engine simulation and test data which show good agreement. Based on a representative transient maneuver typical transient overshoots in fuel flow and turbine gas temperature and changes in component stability margins can be shown. With the use of this model in the performance synthesis the transient engine performance can be predicted more accurate than currently in the engine development program.Copyright © 2004 by ASME
TL;DR: A detailed investigation on the effect of squealer geometries on the blade tip leakage flow and associated heat transfer is presented for a scaled up high pressure turbine blade in a low-speed wind tunnel facility as discussed by the authors.
Abstract: A detailed investigation on the effect of squealer geometries on the blade tip leakage flow and associated heat transfer is presented for a scaled up high pressure turbine blade in a low-speed wind tunnel facility. The linear cascade is made of four blades with the two corner blades acting as guides. The tip profile of a first stage rotor blade is used to fabricate the two-dimensional blade. The wind tunnel accommodates an 116° turn for the blade cascade. The mainstream Reynolds number based on the axial chord length based off cascade exit velocity is 4.83×10 5 . An upstream wake effect is simulated with a spoked wheel wake generator placed upstream of the cascade. A turbulence grid placed even farther upstream generates a free-stream turbulence of 4.8%. The center blade has a tip clearance gap of 1.56% with respect to the blade span. Static pressure measurements are obtained off the blade surface and the shroud
01 Jan 2004
TL;DR: In this paper, experimental and computational studies were carried out in order to survey the energetic aspects of forward and backward sweep in axial flow rotors of low aspect ratio blading for incompressible flow.
Abstract: Experimental and computational studies were carried out in order to survey the energetic aspects of forward and backward sweep in axial flow rotors of low aspect ratio blading for incompressible flow It has been pointed out that negative sweep tends to increase the lift, the flow rate and the ideal total pressure rise in the vicinity of the endwalls Just the opposite tendency was experienced for positive sweep The local losses were found to develop according to combined effects of sweep near the endwalls, endwall and tip clearance losses, and profile drag influenced by re-arrangement of the axial velocity profile The forward-swept bladed rotor showed reduced total efficiency compared to the unswept and swept-back bladed rotors This behavior has been explained on the basis of analysis of flow details It has been found that the swept bladings of low aspect ratio tend to retain the performance of the unswept datum rotor even in absence of sweep correctionCopyright © 2004 by ASME
11 Jul 2004
TL;DR: In this article, a microwave-based sensor designed to operate in temperatures up to 2500°F with a resolution of 0.2 mils and bandwidth up to 25 MHz is presented.
Abstract: Active clearance control has been used to improve efficiency and performance in commercial aircraft engines. Technologies implemented to date rely on compressor bleed air to control clearances based on open loop control and the flight regime (takeoff, cruise, etc.). Open loop systems necessitate a wide safety margin to accommodate the uncertainty in system models that are used to drive such systems. Implementing a feedback control system to optimize clearance has been difficult due to the lack of survivability of clearance measurement technology. Current technologies such as eddy current, capacitive, and laser sensors have been effectively used in laboratory environments but lack the robustness and reliability necessary for long-term use at high engine temperatures. This paper describes a microwave-based sensor designed to operate in temperatures up to 2500°F with a resolution of 0.2 mils and bandwidth up to 25 MHz. The sensor can effectively operate in dirty environments and has the ability to see through oil, combustion products, and other common contaminants. Performance data on the sensor from spin pit testing at 1100°F will be presented to show the viability of the sensor for use in active clearance control.
01 Jan 2004
TL;DR: In this paper, the authors investigate dominant performance limiting mechanisms for micro-scale, high-speed compressor impellers with diameter in the range of 5mm to 10mm and peripheral speed ∼ 500 ms−1.
Abstract: A study has been conducted, using steady three-dimensional Reynolds-averaged Navier-Stokes simulations (FLUENT) to investigate dominant performance limiting mechanisms for micro-scale, high-speed compressor impellers with diameter in the range of 5mm to 10mm and peripheral speed ∼ 500 ms−1 . Heat transfer to impeller flow (hence non-adiabatic in contrast to nearly adiabatic macro-scale centrifugal compressors for aircraft engine application), casing drag, and impeller passage boundary layer loss are identified as micro-scale impeller performance limiting mechanisms. Heat transfer could lead to up to 25 efficiency points penalty, casing drag to about 10–15 points, and passage boundary layer loss to another 10 points for the investigated micro-impellers. Micro-impeller efficiency of up to 90% is achievable if design is directed at mitigating these performance limiting mechanisms. The effect of heat addition on impeller performance is detrimental and depends on a single non-dimensional parameter (ratio of added heat to inlet stagnation enthalpy). The performance penalty is associated with the physical fact that compression at high temperatures requires more work. Casing drag associated with impeller rotating relative to stationary casing results in a torque on the flow near the casing and impeller blade tip that can be characterized in terms of rotational Reynolds number and ratio of tip clearance to impeller radius. Channel boundary layer loss can be characterized in terms of Reynolds number, geometry (impeller exit-to-inlet diameter ratio, blade angles, chord-to-inlet diameter ratio, average-to inlet span ratio, inlet diameter-to-inlet span ratio), and exit-to-inlet temperature ratio related to work input (rotor geometry and speed). A physics-based model is developed for quantifying each of these performance-limiting processes, given the key design parameters. The results from the models are in accord with CFD (FLUENT) data. Implications on impeller design are discussed and design guidelines are formulated. The paper reports a quantitative investigation of micro-turbomachinery performance limiting mechanisms and offers design guidelines based on physical understanding.Copyright © 2004 by ASME
01 Jan 2004
TL;DR: In this paper, a high-speed three-stage axial compressor with inlet guide vanes (IGV) and controlled diffusion airfoils (CDA) was investigated under off-design conditions.
Abstract: This paper deals with unsteady measurements in a highspeed three-stage axial compressor with inlet guide vanes (IGV) and controlled diffusion airfoils (CDA) at off-design conditions. The compressor under consideration exhibits design features of real industrial compressors. The main emphasis is put on the experimental investigation of two operating points at 68% nominal speed where a significant mismatching of the stages occurs. The first operating point is the last stable one near the surge margin whereas the second one represents choke. Probe traverses with a high resolution both in space and time show the significant potential upstream influence of the rotor blades. This effect and the disturbances caused by the convected wakes do strongly influence the unsteady boundary layer behaviour of the stator blades which are detected by glue-on hot-film sensors at different spanwise positions. Dynamic pressure transducers on the casing show that the structure of the rotor tip clearance flow strongly depends on the operating conditions of the compressor. Conclusions can be drawn concerning the consideration of the discussed unsteady effects within the design process of multistage axial compressors with respect to the presented results.Copyright © 2004 by ASME
TL;DR: In this article, the factors that markedly characterise wear are featured, and possible remedies applicable to achieve optimum life are identified, where the concept of accelerated wear test (AWT) is employed to simulate equivalent wear on rotors and bearings; the results obtained are compared with the actual field samples.
01 Jan 2004
TL;DR: In this article, the authors used a series of large-scale simulations to identify the loss mechanisms in a scallop shrouded transonic power generation turbine blade passage at realistic engine conditions.
Abstract: Loss mechanisms in a scallop shrouded transonic power generation turbine blade passage at realistic engine conditions have been identified through a series of large-scale (typically 12 million finite volumes) simulations. All simulations are run with second-order discretization and viscous sublayer resolution, and they include the effects of viscous dissipation. The mesh (y+ near unity on all surfaces) is highly refined in the tip clearance region, casing recesses, and shroud region in order to fully capture complex interdependent flow physics and the associated losses. Aerodynamic losses, in order of their relative importance, are a result of the following: separation around the tip, recesses, and shroud; tip vortex creation; downstream mixing losses, localized shocks on the airfoil; and the passage vortex emanating from under the shroud. A number of helical lateral flows were established near the upper shroud surfaces as a result of lateral pressure gradients on the scalloped shroud. It was found that the tip leakage and passage losses increased approximately linearly with increasing tip clearance, both with and without the effect of the relative casing motion. For each tip clearance studied, scrubbing slightly reduced the tip leakage, but the overall production of entropy was increased by more than 50%. Also the overall passage mass flow rate, for a given inlet total pressure to exit static pressure ratio, increased almost linearly with increasing tip clearance. In addition, it was also found that there was slight positive and negative lift on the shroud, depending on the tip clearance. At the lowest tip clearance of 20 mils there was a negative lift on the shroud. In the 200-mil tip clearance case there was a positive lift on the shroud. The relative motion of the casing contributed positively to the lift at every tip clearance, affecting more at the lowest tip clearance where the casing is closest to the blade tip. Lastly, it was found that the computed entropy generation for the stationary 80-mils case using the SKE turbulence model was close to that of the 80-mils scrubbing case using the RKE turbulence model. In light of the proposed mechanisms and their relative contributions, suggested design considerations are posed.© 2004 ASME