Conditioning of Leakage Flows in Gas Turbine Rotor-Stator Cavities

2020 ◽  
Author(s):  
P. W. Darby ◽  
A. W. Mesny ◽  
G. De Cosmo ◽  
M. Carnevale ◽  
G. D. Lock ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Outer Part ◽  
Engine Performance ◽  
Turbine Rotor ◽  
Flow Visualisation ◽  
Leakage Flow ◽  
Toroidal Vortex ◽  
Performance Measurements ◽  
Operating Cycle ◽  
Leakage Flows

Abstract Ingress is the penetration of hot mainstream fluid into the cavity formed between the turbine disc (rotor) and its adjacent casing (stator). Gas turbine engine designers use rim seals fitted at the periphery of the discs and a superposed sealant flow — typically fed through the bore of the stator — is used to reduce, or in the limit prevent, ingress. Parasitic leakage enters the cavity through pathways created between mating interfaces of engine components. Owing to the aggressive thermal and centrifugal loading experienced during the turbine operating cycle, the degree of leakage and its effect on ingress are difficult to predict. This paper considers the potential for leakage flows to be conditioned in order to minimise their parasitic effect on disc cooling, and ultimately engine, performance. Measurements of static and total pressure, swirl and species concentration were used to assess the performance of a simple axial clearance rim-seal over a range of non-dimensional leakage flow-rates. A computational model was used to provide flow visualisation to support the interpretation of flow structures derived from the experiments. Data is presented to investigate the effects of swirling the leakage flow in accordance with, and counter to, the disc rotation. The injected momentum from the leakage created a toroidal vortex in the outer part of the cavity. Co-swirl was found to improve the sealing effectiveness by up to 15% compared to the axially-introduced baseline and counter-swirled configurations. Varying the momentum of the leakage flow was considered by passing consistent mass-flows through a range of leakage outlet areas. Increasing the momentum was seen to increase the influence of the toroidal vortex on the flow structure in the cavity, which in turn influenced the sealing effectiveness.

Conditioning of Leakage Flows in Gas Turbine Rotor-Stator Cavities

2020 ◽  
Author(s):  
Peter Darby ◽  
Alex Mesny ◽  
Giove De Cosmo ◽  
Mauro Carnevale ◽  
Gary Lock ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Outer Part ◽  
Engine Performance ◽  
Turbine Rotor ◽  
Flow Visualisation ◽  
Leakage Flow ◽  
Performance Measurements ◽  
Operating Cycle ◽  
Turbine Engine ◽  
Leakage Flows

Abstract Ingress is the penetration of hot mainstream fluid into the cavity formed between the turbine disc (rotor) and its adjacent casing (stator). Gas turbine engine designers use rim seals fitted at the periphery of the discs and a superposed sealant flow - typically fed through the bore of the stator - is used to reduce, or in the limit prevent, ingress. Parasitic leakage enters the cavity through pathways created between mating interfaces of engine components. Owing to the aggressive thermal and centrifugal loading experienced during the turbine operating cycle, the degree of leakage and its effect on ingress are difficult to predict. This paper considers the potential for leakage flows to be conditioned in order to minimise their parasitic effect on disc cooling, and ultimately engine, performance. Measurements of static and total pressure, swirl and species concentration were used to assess the performance of a simple axial clearance rim-seal over a range of non-dimensional leakage flow-rates. A computational model was used to provide flow visualisation to support the interpretation of flow structures derived from the experiments. Data is presented to investigate the effects of swirling the leakage flow in accordance with, and counter to, the disc rotation. The injected momentum from the leakage created a toroidal vortex in the outer part of the cavity. Co-swirl was found to improve the sealing effectiveness by up to 15...abridged

Influence of Honeycomb Rubbing on the Labyrinth Seal Performance

2016 ◽  
Vol 139 (1) ◽  
Author(s):  
Daniel Frączek ◽  
Włodzimierz Wróblewski ◽  
Krzysztof Bochon
Keyword(s):  
Aircraft Engine ◽  
Labyrinth Seal ◽  
Turbine Rotor ◽  
Geometrical Model ◽  
Computational Effort ◽  
Radial Clearance ◽  
Leakage Flow ◽  
Leakage Flows ◽  
Full Structure ◽  
Sst Turbulence Model

The aircraft engine operates in various conditions. In consequence, the design of seals must take account of the seal clearance changes and the risk of rubbing. A small radial clearance of the rotor tip seal leads to the honeycomb rubbing in take-off conditions, and the leakage flow may increase in cruise conditions. The aim of this study is to compare two honeycomb seal configurations of the low-pressure gas turbine rotor. In the first configuration, the clearance is small and rubbing occurs. In the second,—the fins of the seal are shorter to eliminate rubbing. It is assumed that the real clearance in both configurations is the same. A study of the honeycomb geometrical model is performed to reduce the computational effort. The problem is investigated numerically using the RANS equations and the two-equation k–ω SST turbulence model. The honeycomb full structure is taken into consideration to show details of the fluid flow. Main parameters of the clearance and leakage flows are compared and discussed for the rotor different axial positions. An assessment of the leakage flow through the seal variants could support the design process.

Conditioning of Leakage Flows in Gas Turbine Rotor-Stator Cavities

2021 ◽  
Author(s):  
Peter Darby ◽  
Alex Mesny ◽  
Giove De Cosmo ◽  
Mauro Carnevale ◽  
Gary Lock ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Turbine Rotor ◽  
Leakage Flows

Flow Migration and Mixing in a High Pressure Turbine Rotor Passage With Axisymmetrically Contoured Endwall and Leakage Flow

2015 ◽  
Author(s):  
Reema Saxena ◽  
Arya Ayaskanta ◽  
Terrence W. Simon ◽  
Hee-Koo Moon ◽  
Luzeng J. Zhang
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Film Cooling ◽  
Thermal Loading ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Leakage Flow ◽  
Flow Rates ◽  
Flow Mixing ◽  
Cooling Effects

The flow field in the passage of a high pressure gas turbine is quite complex, involving strong secondary flows, transverse pressure gradients and strong streamwise acceleration. This complexity may have an adverse effect on cooling of the hub endwall, which is subjected to high thermal loading due to the flat combustor exit temperature profile of modern low-NOx systems. Therefore, given material limitations, better cooling management techniques that can be included with certainty in new gas turbine designs are needed. In the present study, film cooling has been investigated experimentally in a stationary linear cascade. The flow is representative of a high pressure gas turbine rotor with combustor liner coolant introduced to the approach flow. Focus is on the endwall axisymmetric contouring and the cooling effect of leakage flow bled from the compressor through the stator-rotor disc cavity. Two endwall contours, ‘shark nose’ (gradual slope over a larger distance) and ‘dolphin nose’ (steep slope over a shorter distance), are considered and comparison is made under conditions of three mass flow rates (MFR) of leakage, 0.5%, 1.0% and 1.5% of the approach flow rate. The performance of both endwall contours is compared at different streamwise locations in terms of adiabatic effectiveness values over the endwall. This study gives enhanced insight into the physics of coolant flow mixing, migration and subsequent coverage over the endwall. The results show the cooling effects of the contoured shapes over a range of leakage flow rates in the strong secondary flow environment. It is found that the leakage flow plays a crucial role in enhancing coolant coverage over the endwall. To add to our knowledge of mixing effects, detailed thermal field data are taken in the leakage flow discharge region. Doing so helps explain the behavior of the flow as it is ejected into the passage and interacts with the mainstream flow.

Implanted Component Faults and Their Effects on Gas Turbine Engine Performance

1992 ◽  
Vol 114 (2) ◽  
pp. 174-179 ◽  
Author(s):  
J. D. MacLeod ◽  
V. Taylor ◽  
J. C. G. Laflamme
Keyword(s):  
Gas Turbine ◽  
National Research Council ◽  
Engine Performance ◽  
Turbine Rotor ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Turbine Rotor Blade ◽  
Department Of National Defence ◽  
Project Objectives

Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effect is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library, which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: (a) first-stage turbine nozzle erosion damage; (b) first-stage turbine rotor blade untwist; (c) compressor seal wear; (d) first and second-stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.

scholarly journals Effect of in-service burnout on the transonic tip leakage flows over flat tip model

2019 ◽  
Vol 234 (5) ◽  
pp. 655-669
Author(s):  
Zainab Saleh ◽  
Eldad J Avital ◽  
Theodosios Korakianitis
Keyword(s):  
Gas Turbine ◽  
High Speed ◽  
Turbine Blades ◽  
Leakage Flow ◽  
Aerodynamic Loss ◽  
Flow Acceleration ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Operational Life ◽  
Tip Model

Un-shrouded turbine blades are more common than shrouded ones in gas turbine aero-engines since they reduce the weight and avoid the centrifugal loading caused by the blades’ shrouds. Despite these important advantages, the absence of the shroud leads to leakage flows across the tip gap and exposes the blade tip to high thermal load and thermal damages. In addition, the leakage flows can contribute up to 30% of the aerodynamic loss in a turbine stage. In this study, the effect of in-service burnout is explored using a fundamental flat tip model of a high-pressure gas turbine blade. This investigation is carried out both experimentally in a transonic wind tunnel and computationally using the Reynolds Averaged Navier-stokes approach at high-speed conditions. It is found that exposing the tip to the in-service burnout effect changes the leakage flow behaviour significantly when compared with the tip with sharp edges (i.e. the tip at the start of its operational life). Different flow acceleration, flow structure and shockwave pattern and interactions are captured for the round-edge flat tip (i.e. the tip exposed to in-service burnout). The effective tip gap is found to be much larger for the round-edge flat tip allowing more leakage flow into the tip gap which results into higher tip leakage losses in comparison to the sharp-edge tip. Experimental and computational flow visualisations, surface pressure distributions and discharge coefficient are given and analysed for several pressure ratios over the tip gap.

Implanted Component Faults and Their Effects on Gas Turbine Engine Performance

1991 ◽  
Author(s):  
J. D. MacLeod ◽  
V. Taylor ◽  
J. C. G. Laflamme
Keyword(s):  
Gas Turbine ◽  
National Research Council ◽  
Engine Performance ◽  
Turbine Rotor ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Turbine Rotor Blade ◽  
Department Of National Defence ◽  
Project Objectives

Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effort is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: a) 1st stage turbine nozzle erosion damage, b) 1st stage turbine rotor blade untwist, c) compressor seal wear, d) 1st and 2nd stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.

Integrated Lifing Analysis Tool for Gas Turbine Components

2000 ◽  
Author(s):  
Tiedo Tinga ◽  
Wilfried P. J. Visser ◽  
Wim B. de Wolf ◽  
Michael J. Broomhead
Keyword(s):  
Gas Turbine ◽  
Mechanical Load ◽  
Engine Performance ◽  
Turbine Rotor ◽  
System Model ◽  
Analysis Tool ◽  
Turbine Rotor Blade ◽  
Turbine Component ◽  
Engine System ◽  
Life Consumption

A method to predict gas turbine component life based on analysis of engine performance is presented. Engine performance history is obtained from in-flight monitored engine parameters and flight conditions and downloaded for processing by a tool integrating a number of software tools and models. These subsequently include a comprehensive thermodynamical engine system model, heat transfer, thermal and mechanical load models, and finally, a life consumption model. Thermal and mechanical load distributions in the component as well as component life can be predicted. At this stage, the overall life prediction inaccuracy of the tool is dominated by the relatively high inaccuracy of the lifing model, and therefore, component life can only be predicted relative to a reference life. The tool is demonstrated with an analysis of the F100-PW-220 engine 3rd stage turbine rotor blade life consumption during a recorded RNLAF F-16 mission. Using the engine system model with a detailed control system, deterioration effects on engine performance were analyzed and the effect of engine deterioration on blade life consumption rate was determined. The tool has significant potential to enhance on-condition maintenance and optimize aircraft operational use.

Design, Development and Application of Vehicle Gas Turbine Engines

1966 ◽  
Vol 181 (1) ◽  
pp. 149-172
Author(s):  
P. A. Phillips ◽  
Peter Spear
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Low Cost ◽  
Engine Performance ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Design Development ◽  
Turbine Engine ◽  
Wide Range ◽  
Cycle Temperature

After briefly summarizing worldwide automotive gas turbine activity, the paper analyses the power plant requirements of a wide range of vehicle applications in order to formulate the design criteria for acceptable vehicle gas turbines. Ample data are available on the thermodynamic merits of various gas turbine cycles; however, the low cost of its piston engine competitor tends to eliminate all but the simplest cycles from vehicle gas turbine considerations. In order to improve the part load fuel economy, some complexity is inevitable, but this is limited to the addition of a glass ceramic regenerator in the 150 b.h.p. engine which is described in some detail. The alternative further complications necessary to achieve satisfactory vehicle response at various power/weight ratios are examined. Further improvement in engine performance will come by increasing the maximum cycle temperature. This can be achieved at lower cost by the extension of the use of ceramics. The paper is intended to stimulate the design application of the gas turbine engine.

Flow Structures in the Tip Region for a Transonic Compressor Rotor

2013 ◽  
Vol 135 (3) ◽  
Author(s):  
Juan Du ◽  
Feng Lin ◽  
Jingyi Chen ◽  
Chaoqun Nie ◽  
Christoph Biela
Keyword(s):  
Numerical Simulations ◽  
Reference Frame ◽  
Three Dimensional ◽  
Flow Structures ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Three Dimensional Flow ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Tip Leakage Flows

Numerical simulations are carried out to investigate flow structures in the tip region for an axial transonic rotor, with careful comparisons with the experimental results. The calculated performance curve and two-dimensional (2D) flow structures observed at casing, such as the shock wave, the expansion wave around the leading edge, and the tip leakage flow at peak efficiency and near-stall points, are all captured by simulation results, which agree with the experimental data well. An in-depth analysis of three-dimensional flow structures reveals three features: (1) there exists an interface between the incoming main flow and the tip leakage flow, (2) in this rotor the tip leakage flows along the blade chord can be divided into at least two parts according to the blade loading distribution, and (3) each part plays a different role on the stall inception mechanism in the leakage flow dominated region. A model of three-dimensional flow structures of tip leakage flow is thus proposed accordingly. In the second half of this paper, the unsteady features of the tip leakage flows, which emerge at the operating points close to stall, are presented and validated with experiment observations. The numerical results in the rotor relative reference frame are first converted to the casing absolute reference frame before compared with the measurements in experiments. It is found that the main frequency components of simulation at absolute reference frame match well with those measured in the experiments. The mechanism of the unsteadiness and its significance to stability enhancement design are then discussed based on the details of the flow field obtained through numerical simulations.

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