Effect of Non-Uniform Inlet Conditions on Endwall Secondary Flows

2002 ◽  
Author(s):  
K. S. Hermanson ◽  
K. A. Thole
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Total Pressure ◽  
Large Scale ◽  
Secondary Flows ◽  
Turbine Engine ◽  
Thermal Fields ◽  
Pressure Profiles ◽  
Stator Vane ◽  
Inlet Conditions

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly non-uniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give non-uniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flow fields in a first stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flow field data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

Effect of Nonuniform Inlet Conditions on Endwall Secondary Flows

2002 ◽  
Vol 124 (4) ◽  
pp. 623-631 ◽  
Author(s):  
K. S. Hermanson ◽  
K. A. Thole
Keyword(s):  
Secondary Flow ◽  
Total Pressure ◽  
Large Scale ◽  
Secondary Flows ◽  
Turbine Engine ◽  
Thermal Fields ◽  
Pressure Profiles ◽  
Stator Vane ◽  
Inlet Conditions ◽  
Computational Predictions

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly nonuniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give nonuniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flowfields in a first-stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flowfield data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

Effects of Slot Bleed Injection Over a Contoured Endwall on Nozzle Guide Vane Cooling Performance: Part I — Flow Field Measurements

2000 ◽  
Author(s):  
Steven W. Burd ◽  
Terrence W. Simon
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Large Scale ◽  
Secondary Flows ◽  
Cooling Flow ◽  
Nozzle Guide ◽  
Aerodynamic Losses

Film cooling and secondary flows are major contributors to aerodynamic losses in turbine passages. This is particularly true in low aspect ratio nozzle guide vanes where secondary flows can occupy a large portion of the passage flow field. To reduce losses, advanced cooling concepts and secondary flow control techniques must be considered. To this end, combustor bleed cooling flows introduced through an inclined slot upstream of the airfoils in a nozzle passage were experimentally investigated. Testing was performed in a large-scale, high-pressure turbine nozzle cascade comprised of three airfoils between one contoured and one flat endwall. Flow was delivered to this cascade with high-level (∼9%), large-scale turbulence at a Reynolds number based on inlet velocity and true chord length of 350,000. Combustor bleed cooling flow was injected through the contoured endwall upstream of the contouring at bleed-to-core mass flow rate ratios ranging from 0 to 6%. Measurements with triple-sensor, hot-film anemometry characterize the flow field distributions within the cascade. Total and static pressure measurements document aerodynamic losses. The influences of bleed mass flow rate on flow field mean streamwise and cross-stream velocities, turbulence distributions, and aerodynamic losses are discussed. Secondary flow features are also described through these measurements. Notably, this study shows that combustor bleed cooling flow imposes no aerodynamic penalty. This is atypical of schemes where coolant is introduced within the passage for the purpose of endwall cooling. Also, instead of being adversely affected by secondary flows, this type of cooling is able to reduce secondary flow effects.

Combustor Turbine Interface Studies: Part 2 — Flow and Thermal Field Measurements

2002 ◽  
Author(s):  
W. F. Colban ◽  
A. T. Lethander ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Field Measurements ◽  
Pressure Profile ◽  
Stagnation Line ◽  
Turbine Engine ◽  
Turbine Vane ◽  
Cooling Methods

Most turbine inlet flows resulting from the combustor exit are non-uniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identifed from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingment locations along the stagnation line.

Combustor Turbine Interface Studies—Part 2: Flow and Thermal Field Measurements

2003 ◽  
Vol 125 (2) ◽  
pp. 203-209 ◽  
Author(s):  
W. F. Colban ◽  
A. T. Lethander ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Field Measurements ◽  
Pressure Profile ◽  
Stagnation Line ◽  
Turbine Engine ◽  
Turbine Vane ◽  
Cooling Methods

Most turbine inlet flows resulting from the combustor exit are nonuniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identified from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingement locations along the stagnation line.

Flow Field Computations of Combustor-Turbine Interactions Relevant to a Gas Turbine Engine

2004 ◽  
Vol 126 (1) ◽  
pp. 122-129 ◽  
Author(s):  
Sarah Stitzel ◽  
Karen A. Thole
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbine ◽  
High Performance ◽  
Computational Study ◽  
Pressure Profile ◽  
Turbine Engines ◽  
Turbine Engine ◽  
Stator Vane ◽  
Secondary Flow Field

The current demands for high-performance gas turbine engines can be reached by raising combustion temperatures to increase power output. High combustion temperatures create a harsh environment that leads to the consideration of the durability of the combustor and turbine sections. This paper presents a computational study of a flow field that is representative of what occurs in a combustor and how that flow field convects through the first downstream stator vane. The results of this study indicate that the development of the secondary flow field in the turbine is highly dependent on the incoming total pressure profile. The endwall heat transfer is also found to depend strongly on the secondary flow field.

Flow Field Computations of Combustor-Turbine Interactions Relevant to a Gas Turbine Engine

2003 ◽  
Author(s):  
Sarah Stitzel ◽  
Karen A. Thole
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbine ◽  
High Performance ◽  
Computational Study ◽  
Pressure Profile ◽  
Turbine Engines ◽  
Turbine Engine ◽  
Stator Vane ◽  
Secondary Flow Field

The current demands for high performance gas turbine engines can be reached by raising combustion temperatures to increase power output. High combustion temperatures create a harsh environment that leads to the consideration of the durability of the combustor and turbine sections. This paper presents a computational study of a flow field that is representative of what occurs in a combustor and how that flow field convects through the first downstream stator vane. The results of this study indicate the development of the secondary flow field in the turbine is highly dependent on the incoming total pressure profile. The endwall heat transfer is found to also depend strongly on the secondary flow field.

scholarly journals Incidence Angle and Pitch-Chord Effects on Secondary Flows Downstream of a Turbine Cascade

1992 ◽  
Author(s):  
A. Perdichizzi ◽  
V. Dossena
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Incidence Angle ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Turbine Cascade ◽  
Oil Flow ◽  
Secondary Velocity ◽  
Secondary Flow Field

This paper describes the results of an experimental investigation of the three-dimensional flow downstream of a linear turbine cascade at off-design conditions. The tests have been carried out for five incidence angles from −60 to +35 degrees, and for three pitch-chord ratios: s/c = 0.58,0.73,0.87. Data include blade pressure distributions, oil flow visualizations, and pressure probe measurements. The secondary flow field has been obtained by traversing a miniature five hole probe in a plane located at 50% of an axial chord downstream of the trailing edge. The distributions of local energy loss coefficients, together with vorticity and secondary velocity plots show in detail how much the secondary flow field is modified both by incidence and cascade solidity variations. The level of secondary vorticity and the intensity of the crossflow at the endwall have been found to be strictly related to the blade loading occurring in the blade entrance region. Heavy changes occur in the spanwise distributions of the pitch averaged loss and of the deviation angle, when incidence or pitch-chord ratio is varied.

scholarly journals The Design of an Improved Endwall Film-Cooling Configuration

1998 ◽  
Author(s):  
S. Friedrichs ◽  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Aerodynamic Loss ◽  
Cooling Flow ◽  
High Static Pressure ◽  
Endwall Film Cooling ◽  
Aerodynamic Losses

The endwall film-cooling cooling configuration investigated by Friedrichs et al. (1996, 1997) had in principle sufficient cooling flow for the endwall, but in practice, the redistribution of this coolant by secondary flows left large endwall areas uncooled. This paper describes the attempt to improve upon this datum cooling configuration by redistributing the available coolant to provide a better coolant coverage on the endwall surface, whilst keeping the associated aerodynamic losses small. The design of the new, improved cooling configuration was based on the understanding of endwall film-cooling described by Friedrichs et al. (1996, 1997). Computational fluid dynamics were used to predict the basic flow and pressure field without coolant ejection. Using this as a basis, the above described understanding was used to place cooling holes so that they would provide the necessary cooling coverage at minimal aerodynamic penalty. The simple analytical modelling developed in Friedrichs et al. (1997) was then used to check that the coolant consumption and the increase in aerodynamic loss lay within the limits of the design goal. The improved cooling configuration was tested experimentally in a large scale, low speed linear cascade. An analysis of the results shows that the redesign of the cooling configuration has been successful in achieving an improved coolant coverage with lower aerodynamic losses, whilst using the same amount of coolant as in the datum cooling configuration. The improved cooling configuration has reconfirmed conclusions from Friedrichs et al. (1996, 1997); firstly, coolant ejection downstream of the three-dimensional separation lines on the endwall does not change the secondary flow structures; secondly, placement of holes in regions of high static pressure helps reduce the aerodynamic penalties of platform coolant ejection; finally, taking account of secondary flow can improve the design of endwall film-cooling configurations.

Second Vane Total Pressure Loss Due to Endwall Iceform Contouring

2008 ◽  
Author(s):  
Ronald S. LaFleur
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Drag ◽  
Pressure Coefficient ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Shape ◽  
Exit Plane

The iceformation design method generates an endwall contour, altering the secondary flows that produce elevated endwall heat transfer load and total pressure losses. Iceformation is an analog to regions of metal melting where a hot fluid alters the isothermal surface shape of a part as it is maintained by a cooling fluid. The passage flow, heat transfer and geometry evolve together under the constraints of flow and thermal boundary conditions. The iceformation concept is not media dependent and can be used in analogous flows and materials to evolve novel boundary shapes. In the past, this method has been shown to reduce aerodynamic drag and total pressure loss in flows such as diffusers and cylinder/endwall junctures. A prior paper [1] showed that the Reynolds number matched iceform geometry had a 24% lower average endwall heat transfer than the rotationally symmetric endwall geometry of the Energy Efficiency Engine (E3). Comparisons were made between three endwall geometries: the ‘iceform’, the ‘E3’ and the ‘flat’ as a limiting case of the endwall design space. This paper adds to the iceformation design record by reporting the endwall aerodynamic performances. Second vane exit flow velocities and pressures were measured using an automated 2-D traverse of a 1.2 mm diameter five-hole probe. Exit plane maps for the three endwall geometries are presented showing the details of the total pressure coefficient contours and the velocity vectors. The formation of secondary flow vortices is shown in the exit plane and this results in an impact on exit plane total pressure loss distribution, off-design over- and under-turning of the exit flow. The exit plane contours are integrated to form overall measures of the total pressure loss. Relative to the E3 endwall, the iceform endwall has a slightly higher total pressure loss attributed to higher dissipation of the secondary flow within the passage. The iceform endwall has a closer-to-design exit flow pattern than the E3 endwall.

Design Optimization of Profiled Endwall With Consideration of Cooling and Rim Seal Flow Effects

2016 ◽  
Author(s):  
Huimin Tang ◽  
Shuaiqiang Liu ◽  
Hualing Luo
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Secondary Flow ◽  
Three Dimensional ◽  
Great Influence ◽  
Optimization Procedure ◽  
Secondary Flows ◽  
High Pressure Turbine ◽  
Operation Condition ◽  
Turbine Performance

Profiled endwall is an effective method to improve aerodynamic performance of turbine. This approach has been widely studied in the past decade on many engines. When automatic design optimisation is considered, most of the researches are usually based on the assumption of a simplified simulation model without considering cooling and rim seal flows. However, many researchers find out that some of the benefits achieved by optimization procedure are lost when applying the high-fidelity geometry configuration. Previously, an optimization procedure has been implemented by integrating the in-house geometry manipulator, a commercial three-dimensional CFD flow solver and the optimization driver, IsightTM. This optimization procedure has been executed [12] to design profiled endwalls for a turbine cascade and a one-and-half stage axial turbine. Improvements of the turbine performance have been achieved. As the profiled endwall is applied to a high pressure turbine, the problems of cooling and rim seal flows should be addressed. In this work, the effects of rim seal flow and cooling on the flow field of two-stage high pressure turbine have been presented. Three optimization runs are performed to design the profiled endwall of Rotor-One with different optimization model to consider the effects of rim flow and cooling separately. It is found that the rim seal flow has a significant impact on the flow field. The cooling is able to change the operation condition greatly, but barely affects the secondary flow in the turbine. The influences of the profiled endwalls on the flow field in turbine and cavities have been analyzed in detail. A significant reduction of secondary flows and corresponding increase of performance are achieved when taking account of the rim flows into the optimization. The traditional optimization mechanism of profiled endwall is to reduce the cross passage gradient, which has great influence on the strength of the secondary flow. However, with considering the rim seal flows, the profiled endwall improves the turbine performance mainly by controlling the path of rim seal flow. Then the optimization procedure with consideration of rim seal flow has also been applied to the design of the profiled endwall for Stator Two.

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