Aerodynamic performance of an un-located high-pressure turbine rotor

2017 ◽  
Vol 121 (1242) ◽  
pp. 1200-1215 ◽  
Author(s):  
L. Pawsey ◽  
D. J. Rajendran ◽  
V. Pachidis
Keyword(s):  
High Pressure ◽  
Turbine Rotor ◽  
Expansion Ratio ◽  
Leakage Flow ◽  
Flow Properties ◽  
High Pressure Turbine ◽  
Air System ◽  
Secondary Air ◽  
Shaft Failure ◽  
Downstream Movement

ABSTRACTThe rotor sub-assembly of the high-pressure turbine of a modern turbofan engine is typically free to move downstream because of the force imbalance acting on the disc and blades following an un-located shaft failure. This downstream movement results in a change in the geometry of the rotor blade, tip seals and rim/platform seals because of the interaction of the rotor sub-assembly with the downstream vane sub-assembly. Additionally, there is a change in the leakage flow properties, which mix with the main flow because of the change in engine behaviour and secondary air system dynamics. In the present work, the changes in geometry following the downstream movement of the turbine, are obtained from a validated friction model and structural LS-DYNA simulations. Changes in leakage flow properties are obtained from a transient network source-sink secondary air system model. Three-dimensional Reynolds-averaged Navier-Stokes simulations are used to evaluate the aerodynamic effect from the inclusion of the leakage flows, tipseal domains, and downstream movement of the rotor for three displacement configurations (i.e. 0, 10 and 15 mm) with appropriate changes in geometry and leakage flow conditions. It is observed from the results that there is a significant reduction in the expansion ratio, torque and power produced by the turbine with the downstream movement of the rotor because of changes in the flow behaviour for the different configurations. These changes in turbine performance parameters are necessary to accurately predict the terminal speed of the rotor using an engine thermodynamic model. Further, it is to be noted that such reductions in turbine rotor torque will result in a reduction of the terminal speed attained by the rotor during an un-located shaft failure. Therefore the terminal speed of the rotor can be controlled by introducing design features that will result in the rapid rearward displacement of the turbine rotor.

Aerodynamic Performance of an Unlocated High Pressure Turbine Rotor With Worn Tip Seal Fins

2017 ◽  
Author(s):  
Lucas Pawsey ◽  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
High Pressure ◽  
Rotor Blade ◽  
Turbine Rotor ◽  
Aerodynamic Performance ◽  
Axial Displacement ◽  
High Pressure Turbine ◽  
Design Modification ◽  
Turbine Casing ◽  
Secondary Air ◽  
Shaft Failure

An unlocated shaft failure in the high pressure turbine spool of an engine may result in a complex orbiting motion along with rearward axial displacement of the high pressure turbine rotor sub-assembly. This is due to the action of resultant forces and limitations imposed by constraints such as the bearings and turbine casing. Such motion of the rotor following an unlocated shaft failure, results in the development of multiple contacts between the components of the rotor sub-assembly, the turbine casing, and the downstream stator casing. Typically, in the case of shrouded rotor blades, the tip region is in the form of a seal with radial protrusions called ‘fins’ between the rotor blade and the turbine casing. The contact between the rotor blade and the turbine casing will therefore result in excessive wear of the tip seal fins, resulting in changes in the geometry of the tip seal domain that affects the characteristics of the tip leakage vortex. The rotor sub-assembly with worn seals may also be axially displaced rearwards, and consequent to this displacement, changes in the geometry of the rotor blade may occur because of the contact between the rotor sub-assembly and the downstream stator casing. An integrated approach of structural analyses, secondary air system dynamics, and 3D CFD is adopted in the present study to quantify the effect of the tip seal damage and axial displacement on the aerodynamic performance of the turbine stage. The resultant geometry after wearing down of the fins in the tip seal, and rearward axial displacement of the rotor sub-assembly is obtained from LS-DYNA simulations. 3D RANS analyses are carried out to quantify the aerodynamic performance of the turbine with worn fins in the tip seal at three different axial displacement locations i.e. 0 mm, 10 mm and 15 mm. The turbine performance parameters are then compared with equivalent cases in which the fins in the tip seal are intact for the same turbine axial displacement locations. From this study it is noted that the wearing of tip seal fins results in reduced turbine torque, power output and efficiency, consequent to changes in the flow behaviour in the turbine passages. The reduction in turbine torque will result in the reduction of the terminal speed of the rotor during an unlocated shaft failure. Therefore, a design modification that can lead to rapid wearing of the fins in the tip seal after an unlocated shaft failure holds promise for the management of a potential over-speed event.

Prediction of Main/Secondary-Air System Flow Interaction in a High-Pressure Turbine

2005 ◽  
Vol 21 (1) ◽  
pp. 158-166
Author(s):  
Roger L. Davis ◽  
Juan J. Alonso ◽  
Jixian Yao ◽  
Roger Paolillo ◽  
Om P. Sharma
Keyword(s):  
High Pressure ◽  
High Pressure Turbine ◽  
Flow Interaction ◽  
System Flow ◽  
Air System ◽  
Secondary Air

High Pressure Turbine Rotor Stage Endwall Cooling and Passage Thermal Fields as Affected by Leakage Flow and Combustor Liner Coolant Streams

2013 ◽  
Author(s):  
Arya Ayaskanta ◽  
Terrence W. Simon ◽  
Ryan Erickson ◽  
Hee Koo Moon ◽  
Luzeng Zhang
Keyword(s):  
High Pressure ◽  
Turbine Rotor ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
Flow Temperature ◽  
Thermal Fields ◽  
Adiabatic Effectiveness ◽  
Combustor Liner ◽  
Endwall Cooling ◽  
Insight Into

The effects of an engine-representative combustor exit temperature profile and different disc cavity leakage flow rates on endwall adiabatic effectiveness distributions and passage temperature fields in a high pressure turbine rotor stage of a gas turbine are experimentally documented. The measurements are made on a stationary linear blade row cascade with an axisymmetrically-contoured endwall of modern engine geometry and with engine-representative approach flow thermal and fluid mechanics characteristics. The measurements give insight into mixing of coolant emerging as leakage flow and combustor liner coolant mix with hot core gases ahead of the airfoil row. Reported results are thermal fields in the passage, adiabatic wall temperatures and adiabatic effectiveness values in using an engine-representative approach flow temperature profile and with approach flow temperature profiles with 1) no coolant in the approach flow (flat profile) and 2) coolant only within 10% of the span (approach flow profile with a thin thermal boundary layer).The results give insight into mixing between the leakage flow and the mainstream passage flow and its effects on endwall cooling. The results demonstrate that for the conditions studied; much of the endwall cooling is contributed by the coolant in the approach flow. This is an important result that has previously not been well documented.

Tip Clearance Effects on Inlet Hot Streaks Migration Characteristics in High Pressure Stage of a Vaneless Counter-Rotating Turbine

2008 ◽  
Author(s):  
Qingjun Zhao ◽  
Jianyi Du ◽  
Huishe Wang ◽  
Xiaolu Zhao ◽  
Jianzhong Xu
Keyword(s):  
High Pressure ◽  
Turbine Rotor ◽  
Tip Clearance ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Pressure Surface ◽  
Low Pressure Turbine ◽  
Turbine Stator ◽  
Hot Streak

In this paper, three-dimensional multiblade row unsteady Navier-Stokes simulations at a hot streak temperature ratio of 2.0 have been performed to reveal the effects of rotor tip clearance on the inlet hot streak migration characteristics in high pressure stage of a Vaneless Counter-Rotating Turbine. The hot streak is circular in shape with a diameter equal to 25% of the high pressure turbine stator span. The hot streak center is located at 50% of the span and the leading edge of the high pressure turbine stator. The tip clearance size studied in this paper is 2.0mm (2.594% high pressure turbine rotor height). The numerical results indicate that the hot streak mixes with the high pressure turbine stator wake and convects towards the high pressure turbine rotor blade surface. Most of hotter fluid migrates to the pressure surface of the high pressure turbine rotor. Only a few of hotter fluid rounds the leading edge of the high pressure turbine rotor and migrates to the suction surface. The migration characteristics of the hot streak in the high pressure turbine rotor are dominated by the combined effects of secondary flow, buoyancy and leakage flow in the rotor tip clearance. The leakage flow trends to drive the hotter fluid towards the blade tip on the pressure surface and to the hub on the suction surface. Under the effect of the leakage flow, even partial hotter fluid near the pressure surface is also driven to the rotor suction surface through the tip clearance. Compared with the case without rotor tip clearance, the heat load of the high pressure turbine rotor is intensified due to the effects of the leakage flow. And the results indicate that the leakage flow effects trend to increase the low pressure turbine rotor inlet temperature at the tip region. The air flow with higher temperature at the tip region of the low pressure turbine rotor inlet will affect the flow and heat transfer characteristics in the downstream low pressure turbine.

Influence of Hot Streak Temperature Ratio on Low Pressure Stage of a Vaneless Counter-Rotating Turbine

2007 ◽  
Author(s):  
Qingjun Zhao ◽  
Fei Tang ◽  
Huishe Wang ◽  
Jianyi Du ◽  
Xiaolu Zhao ◽  
...  
Keyword(s):  
Shock Wave ◽  
High Pressure ◽  
Secondary Flow ◽  
Low Pressure ◽  
Turbine Rotor ◽  
Temperature Ratio ◽  
High Pressure Turbine ◽  
Pressure Stage ◽  
Low Pressure Turbine ◽  
Hot Streak

In order to explore the influence of hot streak temperature ratio on low pressure stage of a Vaneless Counter-Rotating Turbine, three-dimensional multiblade row unsteady Navier-Stokes simulations have been performed. The predicted results show that hot streaks are not mixed out by the time they reach the exit of the high pressure turbine rotor. The separation of colder and hotter fluids is observed at the inlet of the low pressure turbine rotor. After making interactions with the inner-extending shock wave and outer-extending shock wave in the high pressure turbine rotor, the hotter fluid migrates towards the pressure surface of the low pressure turbine rotor, and the most of colder fluid migrates to the suction surface of the low pressure turbine rotor. The migrating characteristics of the hot streaks are predominated by the secondary flow in the low pressure turbine rotor. The effect of buoyancy on the hotter fluid is very weak in the low pressure turbine rotor. The results also indicate that the secondary flow intensifies in the low pressure turbine rotor when the hot streak temperature ratio is increased. The effects of the hot streak temperature ratio on the relative Mach number and the relative flow angle at the inlet of the low pressure turbine rotor are very remarkable. The isentropic efficiency of the Vaneless Counter-Rotating Turbine decreases as the hot streak temperature ratio is increased.

Impact of Individual High-Pressure Turbine Rotor Purge Flows on Turbine Center Frame Aerodynamics

2017 ◽  
Author(s):  
S. Zerobin ◽  
C. Aldrian ◽  
A. Peters ◽  
F. Heitmeir ◽  
E. Göttlich
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Turbine Rotor ◽  
Flow Conditions ◽  
High Pressure Turbine ◽  
Reference Case ◽  
University Of Technology ◽  
Purge Flow ◽  
Generation Mechanisms ◽  
The Impact

This paper presents an experimental study of the impact of individual high-pressure turbine purge flows on the main flow in a downstream turbine center frame duct. Measurements were carried out in a product-representative one and a half stage turbine test setup, installed in the Transonic Test Turbine Facility at Graz University of Technology. The rig allows testing at engine-relevant flow conditions, matching Mach, Reynolds, and Strouhal number at the inlet of the turbine center frame. The reference case features four purge flows differing in flow rate, pressure, and temperature, injected through the hub and tip, forward and aft cavities of the high-pressure turbine rotor. To investigate the impact of each individual cooling flow on the flow evolution in the turbine center frame, the different purge flows were switched off one-by-one while holding the other three purge flow conditions. In total, this approach led to six different test conditions when including the reference case and the case without any purge flow ejection. Detailed measurements were carried out at the turbine center frame duct inlet and outlet for all six conditions and the post-processed results show that switching off one of the rotor case purge flows leads to an improved duct performance. In contrast, the duct exit flow is dominated by high pressure loss regions if the forward rotor hub purge flow is turned off. Without the aft rotor hub purge flow, a reduction in duct pressure loss is determined. The purge flows from the rotor aft cavities are demonstrated to play a particularly important role for the turbine center frame aerodynamic performance. In summary, this paper provides a first-time assessment of the impact of four different purge flows on the flow field and loss generation mechanisms in a state-of-the-art turbine center frame configuration. The outcomes of this work indicate that a high-pressure turbine purge flow reduction generally benefits turbine center frame performance. However, the forward rotor hub purge flow actually stabilizes the flow in the turbine center frame duct and reducing this purge flow can penalize turbine center frame performance. These particular high-pressure turbine/turbine center frame interactions should be taken into account whenever high-pressure turbine purge flow reductions are pursued.

Heat Transfer and Aerodynamics of Over-Shroud Leakage Flows in a High-Pressure Turbine

2009 ◽  
Author(s):  
Knut Lehmann ◽  
Richard Thomas ◽  
Howard Hodson ◽  
Vassilis Stefanis
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Large Scale ◽  
Suction Side ◽  
Surface Flow ◽  
Tip Clearance ◽  
Flow Visualisation ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
High Heat Transfer

An experimental study has been conducted to investigate the distribution of the convective heat transfer on the shroud of a high pressure turbine blade in a large scale rotating rig. A continuous thin heater foil technique has been adapted and implemented on the turbine shroud. Thermochromic Liquid Crystals were employed for the surface temperature measurements to derive the experimental heat transfer data. The heat transfer is presented on the shroud top surfaces and the three fins. The experiments were conducted for a variety of Reynolds numbers and flow coefficients. The effects of different inter-shroud gap sizes and reduced fin tip clearance gaps were also investigated. Details of the shroud flow field were obtained using an advanced Ammonia-Diazo surface flow visualisation technique. CFD predictions are compared with the experimental data and used to aid interpretation. Contour maps of the Nusselt number reveal that regions of highest heat transfer are mostly confined to the suction side of the shroud. Peak values exceed the average by as much as 100 percent. It has been found that the interaction between leakage flow through the inter-shroud gaps and the fin tip leakage jets are responsible for this high heat transfer. The inter-shroud gap leakage flow causes a disruption of the boundary layer on the turbine shroud. Furthermore, the development of the large recirculating shroud cavity vortices is severely altered by this leakage flow.

Multistage Compressor and Turbine Modeling for the Prediction of the Maximum Turbine Speed Resulting From Shaft Breakage

2009 ◽  
Author(s):  
M. Haake ◽  
R. Fiola ◽  
S. Staudacher
Keyword(s):  
High Pressure ◽  
Tip Clearance ◽  
Maximum Speed ◽  
High Pressure Turbine ◽  
High Pressure Compressor ◽  
Failure Event ◽  
Turbine Speed ◽  
Shaft Failure ◽  
Pressure Compressor ◽  
Vane Tip

A mathematical model for the prediction of the maximum speed of a high pressure turbine following a shaft failure event was developed. The model predicts the high pressure compressor and ducting system pre- and post-stall behavior like rotating stall and surge after the shaft breakage. The corresponding time-dependent high pressure turbine inlet conditions are used to calculate the turbine maximum speed, taking into account friction and blade&vane tip clearance variations as a result of the rearward movement of the turbine and destruction of the turbine blading. The compressor and ducting system is modeled by a 1-dimensional, stage-by-stage approach. The approach uses a finite-difference numerical technique to solve the nonlinear system of equations for continuity, momentum and energy including source terms for the compressible flow through inlet ducting, compressor and combustor. The compressor blade forces and shaft work are provided by a set of quasi steady state stage characteristics being valid for pre-stall and post-stall operations. The maximum turbine speed is calculated from a thermodynamic turbine stand-alone model, derived from a performance synthesis program. Friction and blade&vane tip clearance variations are determined iteratively from graphical data depending on the axial rearward movement of the turbine. The compressor and ducting system model was validated in pre-stall and post-stall operation mode with measured high pressure compressor data of a modern 2-shaft engine. The turbine model was validated with measured intermediate pressure shaft failure data of a 3-shaft engine. The shaft failure model was applied on a modern 2-shaft engine. The model was used to carry out a sensitivity study to demonstrate the impact of control system reactions on the resulting maximum high pressure turbine speed following a shaft failure event.

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