The Effect of Rotor Tip Clearance Size Onto the Separated Flow Through a Super-Aggressive S-Shaped Intermediate Turbine Duct Downstream of a Transonic Turbine Stage

2009 ◽  
Author(s):  
A. Marn ◽  
E. Go¨ttlich ◽  
F. Malzacher ◽  
H. P. Pirker
Keyword(s):  
Low Pressure ◽  
Tip Clearance ◽  
Flow Visualisation ◽  
Turbine Stage ◽  
Intermediate Turbine Duct ◽  
Low Pressure Turbine ◽  
Oil Flow ◽  
Blade Tip ◽  
Flow Through ◽  
Aero Engines

The demand of further increased bypass ratio of aero engines will lead to low pressure turbines with larger diameters, which rotate at lower speed. Therefore, it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at a larger diameter without any loss generating separation or flow disturbances. Due to costs and weight this intermediate turbine duct (ITD) has to be as short as possible. This leads to an aggressive (high diffusion) and further to a super-aggressive s-shaped duct geometry. In order to investigate the influence of the blade tip gap size on such a high diffusion duct flow a detailed test arrangement under engine representative conditions is necessary. Therefore, the continuously operating Transonic Test Turbine Facility (TTTF) at Graz University of Technology has been adapted: An super-aggressive intermediate duct is arranged downstream of a transonic HP-turbine stage providing an exit Mach number of about 0.6 and a swirl angle of −15 degrees. A second LP-vane row is located at the end of the duct and represents the counter rotating low pressure turbine at a larger diameter. A following deswirler and a diffuser are the connection to the exhaust casing of the facility. In order to determine the influence of the blade tip gap size on the flow through such a super-aggressive s-shaped turbine duct measurements were conducted with two different tip gap sizes, 1.5% span (0.8 mm) and 2.4% span (1.3 mm). The aerodynamic design of the HP-turbine stage, ITD, LP-vane and the de-swirler was done by MTU Aero engines. In 2007 at ASME Turbo Expo the influence of the rotor clearance size onto the flow through an aggressive ITD was presented. For the present investigation this aggressive duct has been further shortened by 20% (super-aggressive ITD) that the flow at the outer duct wall is fully separated. This paper shows the influence of the rotor tip clearance size onto this separation. The flow through this intermediate turbine duct was investigated by means of five-hole-probes, static pressure taps, boundary layer rakes and oil flow visualisation. The oil flow visualisation showed the existence of vortical structures within the separation where they seem to be imposed by the upstream HP-vanes. This work is part of the EU-project AIDA (Aggressive Intermediate Duct Aerodynamics, Contract: AST3-CT-2003-502836).

The Effect of Rotor Tip Clearance Size onto the Separated Flow Through a Super-Aggressive S-Shaped Intermediate Turbine Duct Downstream of a Transonic Turbine Stage

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
A. Marn ◽  
E. Göttlich ◽  
F. Malzacher ◽  
H. P. Pirker
Keyword(s):  
Low Pressure ◽  
Tip Clearance ◽  
Turbine Stage ◽  
Intermediate Turbine Duct ◽  
Low Pressure Turbine ◽  
Oil Flow ◽  
Oil Flow Visualization ◽  
Blade Tip ◽  
Flow Through ◽  
Aero Engines

The demand for a further increased bypass ratio of aero engines will lead to low pressure turbines with larger diameters, which rotate at a lower speed. Therefore, it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at a larger diameter without any loss generating separation or flow disturbances. Due to costs and weight, this intermediate turbine duct (ITD) has to be as short as possible. This leads to an aggressive (high diffusion) and, furthermore, to a super-aggressive s-shaped duct geometry. In order to investigate the influence of the blade tip gap size on such a high diffusion duct flow a detailed test arrangement under engine representative conditions is necessary. Therefore, the continuously operating Transonic Test Turbine Facility (TTTF) at Graz University of Technology has been adapted: An super-aggressive intermediate duct is arranged downstream of a transonic high pressure (HP)-turbine stage providing an exit Mach number of about 0.6 and a swirl angle of –15 deg. A second low pressure (LP)-vane row is located at the end of the duct and represents the counter-rotating low pressure turbine at a larger diameter. A following deswirler and a diffuser are the connection to the exhaust casing of the facility. In order to determine the influence of the blade tip gap size on the flow through such a super-aggressive s-shaped turbine, duct measurements were conducted with two different tip gap sizes, a 1.5% span (0.8 mm) and a 2.4% span (1.3 mm). The aerodynamic design of the HP-turbine stage, ITD, LP-vane, and the de-swirler was done by MTU Aero engines. In 2007 at the ASME Turbo Expo, the influence of the rotor clearance size onto the flow through an aggressive ITD was presented. For the present investigation, this aggressive duct has been further shortened by 20% (super-aggressive ITD) so that the flow at the outer duct wall is fully separated. This paper shows the influence of the rotor tip clearance size on to this separation. The flow through this intermediate turbine duct was investigated by means of five-hole-probes, static pressure taps, boundary layer rakes, and oil flow visualization. The oil flow visualization showed the existence of vortical structures within the separation where they seem to be imposed by the upstream HP-vanes.

The Influence of Blade Tip Gap Variation on the Flow Through an Aggressive S-Shaped Intermediate Turbine Duct Downstream a Transonic Turbine Stage: Part I — Time-Averaged Results

2007 ◽  
Author(s):  
A. Marn ◽  
E. Go¨ttlich ◽  
R. Pecnik ◽  
F. J. Malzacher ◽  
O. Schennach ◽  
...  
Keyword(s):  
Low Pressure ◽  
Laser Doppler Velocimeter ◽  
Gap Size ◽  
Turbine Stage ◽  
Intermediate Turbine Duct ◽  
Low Pressure Turbine ◽  
High Diffusion ◽  
Blade Tip ◽  
Flow Through ◽  
Aero Engines

The demand of further increased bypass ratio of aero engines will lead to low pressure turbines with larger diameters which rotate at lower speed. Therefore, it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at a larger diameter without any loss generating separation or flow disturbances. Due to costs and weight this intermediate turbine duct has to be as short as possible. This leads to an aggressive (high diffusion) s-shaped duct geometry. To investigate the influence of the blade tip gap size of such a nonseparating high diffusion duct flow a detailed test arrangement under engine representative conditions is necessary. Therefore, the continuously operating Transonic Test Turbine Facility (TTTF) at Graz University of Technology has been adapted: A high diffusion intermediate duct is arranged downstream a HP turbine stage providing an exit Mach number of about 0.6 and a swirl angle of 15 degrees (counter swirl). An LP vane row is located at the end of the duct and represents the counter rotating low pressure turbine at larger diameter. In order to determine the influence of the blade tip gap size on the flow through such an s-shaped turbine duct measurements were conducted with two different tip gap sizes, 1.5% span (0.8mm) and 2.4% span. (1.3mm). The aerodynamic design of the HP vane, the HP turbine, the duct and the LP vane was done by MTU Aero Engines. The investigation was conducted by means of five-hole-probes with thermocouples, boundary layer rakes and static pressure taps at the inner and outer wall along the duct at several circumferential positions. Five-hole-probe measurements were done in five planes within the duct and in two planes downstream of the LP vane. A rough estimation of the duct loss is given at the end of the paper. Part II of this work deals with two-component Laser-Doppler Velocimeter (LDV) measurements at duct inlet directly downstream the HP blade to obtain unsteady information about the inflow. Additionally, oil film visualisation was used to get information about the surface flow at the outer and inner wall of the duct.

The Influence of Blade Tip Gap Variation on the Flow Through an Aggressive S-Shaped Intermediate Turbine Duct Downstream a Transonic Turbine Stage: Part II — Time-Resolved Results and Surface Flow

2007 ◽  
Author(s):  
E. Go¨ttlich ◽  
A. Marn ◽  
R. Pecnik ◽  
F. J. Malzacher ◽  
O. Schennach ◽  
...  
Keyword(s):  
Low Pressure ◽  
Surface Flow ◽  
Gap Size ◽  
Turbine Stage ◽  
Intermediate Turbine Duct ◽  
Low Pressure Turbine ◽  
High Diffusion ◽  
Blade Tip ◽  
Flow Through ◽  
Aero Engines

The demand of further increased bypass ratio of aero engines will lead to low pressure turbines with larger diameters rotating at lower speed. Therefore it is necessary to guide the flow leaving the high pressure turbine to the low pressure turbine at larger diameter without any separation or flow disturbances. Due to costs and weight this intermediate turbine duct has to be as short as possible leading to aggressive (high diffusion) S-shaped duct geometries. To investigate the influence of the blade tip gap size on such a nonseparating high diffusion duct flow a detailed test arrangement under engine representative conditions is necessary. Therefore the continuously operating Transonic Test Turbine Facility (TTTF) at Graz University of Technology has been adapted: An high diffusion intermediate duct is arranged downstream of a HP turbine stage providing an exit Mach number of about 0.6 and a swirl angle of −15 degrees. A LP vane row is located at the end of the duct and represents the counter rotating low pressure turbine at larger diameter. In order to determine the influence of the blade tip gap size on the flow through such an S-shaped turbine duct measurements were performed with two different tip gap sizes, 0.8 mm and 1.3 mm. The aerodynamic design was done by MTU Aero Engines. While Part I describes the investigation by means of five hole probes with thermo couples, boundary layer rakes and static pressure tappings Part II uses Laser-Doppler-Velocimetry (LDV) for measurements at duct inlet directly downstream the HP blades to obtain unsteady information about the inflow and to quantify the differences between the two tip gaps. Additionally oil-film visualization was used to discuss the surface flow at the outer and inner wall of the duct. A comparison with a numerical simulation is also given. This work is part of the EU-project AIDA (Aggressive Intermediate Duct Aerodynamics, Contract: AST3-CT-2003-502836).

Impact of Passive Tip-Injection on Tip-Leakage Flow in Axial Low Pressure Turbine Stage

2015 ◽  
Author(s):  
Pouya Ghaffari ◽  
Reinhard Willinger ◽  
Sabine Bauinger ◽  
Andreas Marn
Keyword(s):  
Aerodynamic Resistance ◽  
Low Pressure ◽  
Leakage Flow ◽  
Turbine Stage ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Low Pressure Turbine ◽  
Blade Tip ◽  
Tip Injection ◽  
The Impact

In addition to geometrical modifications of the blade tip for reducing tip-leakage mass flow rate the method of passive tip-injection serves as an aerodynamic resistance towards the tip-leakage flow. The impact of this method has been investigated thoroughly at unshrouded blades in linear cascades. Furthermore combinations of shrouded blades with passive tip-injection have been investigated analytically as well as via numerical simulations for incompressible flow in linear cascades. The objective of this paper is to consider a real uncooled low pressure turbine stage with shrouded blades and to investigate the effect of passive tip-injection on various operational characteristics. CFD calculations have been carried out in a rotational frame taking into consideration compressible flow and serve for evaluating the method of passive tip-injection in the given turbine stage. Experimental data obtained from the machine without tip-injection serve as boundary conditions for the CFD calculations.

Improvement of Aerodynamic Performance of a Low-Pressure Turbine Stage for Aero-Engines Taking Advantage of Wake-Blade Interaction

2002 ◽  
Vol 2002 (0) ◽  
pp. 103-104
Author(s):  
Kenichi FUNAZAKI ◽  
Toru TAKASHIMA ◽  
Satoru YAMAMOTO
Keyword(s):  
Low Pressure ◽  
Aerodynamic Performance ◽  
Turbine Stage ◽  
Low Pressure Turbine ◽  
Aero Engines

Design of an Intermediate Turbine Duct Downstream of a High Pressure Turbine

2016 ◽  
Author(s):  
Chaoshan Hou ◽  
Hu Wu
Keyword(s):  
High Pressure ◽  
Three Dimensional ◽  
Low Pressure ◽  
Aerodynamic Load ◽  
Duct Flow ◽  
Original Design ◽  
High Pressure Turbine ◽  
Intermediate Turbine Duct ◽  
Low Pressure Turbine ◽  
Inlet Conditions

The flow leaving the high pressure turbine should be guided to the low pressure turbine by an annular diffuser, which is called as the intermediate turbine duct. Flow separation, which would result in secondary flow and cause great flow loss, is easily induced by the negative pressure gradient inside the duct. And such non-uniform flow field would also affect the inlet conditions of the low pressure turbine, resulting in efficiency reduction of low pressure turbine. Highly efficient intermediate turbine duct cannot be designed without considering the effects of the rotating row of the high pressure turbine. A typical turbine model is simulated by commercial computational fluid dynamics method. This model is used to validate the accuracy and reliability of the selected numerical method by comparing the numerical results with the experimental results. An intermediate turbine duct with eight struts has been designed initially downstream of an existing high pressure turbine. On the basis of the original design, the main purpose of this paper is to reduce the net aerodynamic load on the strut surface and thus minimize the overall duct loss. Full three-dimensional inverse method is applied to the redesign of the struts. It is revealed that the duct with new struts after inverse design has an improved performance as compared with the original one.

scholarly journals Numerical Calculation of Flow for Cascade With Tip Clearance

1994 ◽  
Author(s):  
Shimpei Mizuki ◽  
Hoshio Tsujita
Keyword(s):  
Three Dimensional ◽  
Tip Clearance ◽  
Turbine Cascade ◽  
Governing Equations ◽  
Tensor Form ◽  
Pressure Surface ◽  
Rear Part ◽  
Blade Tip ◽  
Flow Through ◽  
Blade Tip Geometry

Three-dimensional incompressible turbulent flow within a linear turbine cascade with tip clearance is analyzed numerically. The governing equations involving the standard k-ε model are solved in the physical component tensor form with a boundary-fitted coordinate system. In the analysis, the blade tip geometry is treated accurately in order to predict the flow through the tip clearance in detail when the blades have large thicknesses. Although the number of grids employed in the present study is not enough because of the limitation of computer storage memory, the computed results show good agreements with the experimental results. Moreover, the results clearly exhibit the locus of minimum pressure on the rear part of the pressure surface at the blade tip.

scholarly journals 3D shape optimisation of a low-pressure turbine stage

2018 ◽  
Vol 3 (1) ◽  
pp. 10-21
Author(s):  
L. Witanowski ◽  
P. Klonowicz ◽  
P. Lampart
Keyword(s):  
Low Pressure ◽  
Shape Optimisation ◽  
Turbine Stage ◽  
3D Shape ◽  
Low Pressure Turbine

Experimental Study on the Effect of Clocking on the Propagation of Inflow Pressure Distortions in a Low Pressure Turbine Stage

2020 ◽  
Author(s):  
L. Simonassi ◽  
M. Zenz ◽  
P. Bruckner ◽  
S. Pramstrahler ◽  
F. Heitmeir ◽  
...  
Keyword(s):  
Experimental Study ◽  
Flow Field ◽  
Total Pressure ◽  
Low Pressure ◽  
Vibration Characteristics ◽  
Test Facility ◽  
Turbine Stage ◽  
Rotor Vibration ◽  
Low Pressure Turbine ◽  
The Impact

Abstract The design of modern aero engines enhances the interaction between components and facilitates the propagation of circumferential distortions of total pressure and temperature. As a consequence, the inlet conditions of a real turbine have significant spatial non-uniformities, which have direct consequences on both its aerodynamic and vibration characteristics. This work presents the results of an experimental study on the effects of different inlet total pressure distortion-stator clocking positions on the propagation of total pressure inflow disturbances through a low pressure turbine stage, with a particular focus on both the aerodynamic and aeroelastic performance. Measurements at a stable engine relevant operating condition and during transient operation were carried out in a one and a half stage subsonic turbine test facility at the Institute of Thermal Turbomachinery and Machine Dynamics at Graz University of Technology. A localised total pressure distortion was generated upstream of the stage in three different azimuthal positions relative to the stator vanes. The locations were chosen in order to align the distortion directly with a vane leading edge, suction side and pressure side. Additionally, a setup with clean inflow was used as reference. Steady and unsteady aerodynamic measurements were taken downstream of the investigated stage by means of a five-hole-probe (5HP) and a fast response aerodynamic pressure probe (FRAPP) respectively. Strain gauges applied on different blades were used in combination with a telemetry system to acquire the rotor vibration data. The aerodynamic interactions between the stator and rotor rows and the circumferential perturbation were studied through the identification of the main structures constituting the flow field. This showed that the steady and unsteady alterations created by the distortion in the flow field lead to modifications of the rotor vibration characteristics. Moreover, the importance of the impact that the pressure distortion azimuthal position has on the LPT stage aerodynamics and vibrations was highlighted.

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