Investigation of Combustor-Turbine-Interaction in a Rotating Cooled Transonic High-Pressure Turbine Test Rig: Part 2 — Numerical Modelling and Simulation

2019 ◽  
Author(s):  
Simon Gövert ◽  
Federica Ferraro ◽  
Alexander Krumme ◽  
Clemens Buske ◽  
Marc Tegeler ◽  
...  
Keyword(s):  
High Pressure ◽  
Flame Stabilization ◽  
Test Facility ◽  
Test Rig ◽  
High Pressure Turbine ◽  
Lean Burn ◽  
Turbine Inlet ◽  
Measurement Results ◽  
Aero Engine ◽  
Inlet Conditions

Abstract Reducing the uncertainties in the prediction of turbine inlet conditions is a crucial aspect to improve aero engine designs and further increase engine efficiencies. To meet constantly stricter emission regulations, lean burn combustion could play a key role for future engine designs. However, these combustion systems are characterized by significant swirl for flame stabilization and reduced cooling air mass flows. As a result, substantial spatial and transient variations of the turbine inlet conditions are encountered. To investigate the effect of the combustor on the high pressure turbine, a rotating cooled transonic high-pressure configuration has been designed and investigated experimentally at the DLR turbine test facility ‘NG-Turb’ in Göttingen, Germany. It is a rotating full annular 1.5 stage turbine configuration which is coupled to a combustor simulator. The combustor simulator is designed to create turbine inlet conditions which are hydrodynamically representative for a lean-burn aero engine. A detailed description of the test rig and its instrumentation as well as a discussion of the measurement results is presented in part I of this paper. Part II focuses on numerical modeling of the test rig to further extend the understanding of the measurement results. Integrated simulations of the configuration including combustor simulator and nozzle guide vanes are performed for leading edge and passage clocking position and the effect on the hot streak migration is discussed. The simulation and experimental results at the combustor-turbine interface are compared showing a good overall agreement. The relevant flow features are correctly predicted in the simulations, proving the suitability of the numerical model for application to integrated combustor-turbine interaction analysis.

Flow Field and Hot Streak Migration Through a High Pressure Cooled Vanes With Representative Lean Burn Combustor Outflow

2018 ◽  
Vol 141 (4) ◽  
Author(s):  
Tommaso Bacci ◽  
Tommaso Lenzi ◽  
Alessio Picchi ◽  
Lorenzo Mazzei ◽  
Bruno Facchini
Keyword(s):  
Experimental Data ◽  
High Pressure ◽  
Flow Field ◽  
Fuel Injection ◽  
Leading Edge ◽  
Flame Stabilization ◽  
Lean Burn ◽  
Aero Engine ◽  
University Of Florence ◽  
Hot Streak

Modern lean burn aero-engine combustors make use of relevant swirl degrees for flame stabilization. Moreover, important temperature distortions are generated, in tangential and radial directions, due to discrete fuel injection and liner cooling flows respectively. At the same time, more efficient devices are employed for liner cooling and a less intense mixing with the mainstream occurs. As a result, aggressive swirl fields, high turbulence intensities, and strong hot streaks are achieved at the turbine inlet. In order to understand combustor-turbine flow field interactions, it is mandatory to collect reliable experimental data at representative flow conditions. While the separated effects of temperature, swirl, and turbulence on the first turbine stage have been widely investigated, reduced experimental data is available when it comes to consider all these factors together.In this perspective, an annular three-sector combustor simulator with fully cooled high pressure vanes has been designed and installed at the THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion cooled liners, and six film cooled high pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central NGV aligned with the central swirler. In order to generate representative conditions, a heated mainstream passes though the axial swirlers of the combustor simulator, while the effusion cooled liners are fed by air at ambient temperature. The resulting flow field exiting from the combustor simulator and approaching the cooled vane can be considered representative of a modern Lean Burn aero engine combustor with swirl angles above ±50 deg, turbulence intensities up to about 28% and maximum-to-minimum temperature ratio of about 1.25. With the final aim of investigating the hot streaks evolution through the cooled high pressure vane, the mean aerothermal field (temperature, pressure, and velocity fields) has been evaluated by means of a five-hole probe equipped with a thermocouple and traversed upstream and downstream of the NGV cascade.

Assessment of RANS Against LES for the Aero-Thermal Behavior of High Pressure Turbine Stages Under Realistic Inlet Conditions

2016 ◽  
Author(s):  
Stefano Vagnoli ◽  
Tom Verstraete ◽  
Charlie Koupper ◽  
Guillaume Bonneau
Keyword(s):  
High Pressure ◽  
Thermal Behavior ◽  
Turbulence Modeling ◽  
Engine Performance ◽  
High Pressure Turbine ◽  
Lean Burn ◽  
Field Development ◽  
Safety Margins ◽  
Inlet Conditions ◽  
Hot Streak

Modern Lean Burn combustors generate a complex field at the High Pressure turbine (HPT) inlet, characterized by non-uniform velocity and temperature distributions, together with very high turbulence levels (up to 25%). For these extreme conditions, classical numerical methods employed for the HPT design, such as Reynolds Averaged Navier Stokes (RANS) simulation, suffer from a lack of validation. This leads to a reduced confidence in predicting the combustor-turbine interactions, which requires to use extra safety margins, to the detriment of the overall engine performance. Within the European FACTOR project, a 360° non reactive combustor simulator and a 1.5 HPT stage are designed to get more insight into the mutual interaction of these two components. A first experimental and numerical campaign has demonstrated the potential of Large Eddy Simulations (LES) to accurately reproduce the turbulent flow field development at the combustor outlet. The aim of the present paper is to exploit the accuracy of LES to validate less time-consuming RANS models in predicting the hot streak migration in the turbine stage. In this sense, LES results are used as a reference to discriminate the different RANS simulations in terms of turbulence modeling and aerothermal predictions. The current investigations clearly indicate that turbulence and hot streak diffusion within the HPT are strongly linked. In this sense, the choice of the RANS turbulence model and the inlet turbulent conditions plays a major role in modeling the thermal behavior for the stator and rotor blades.

Effect of Low-NOX Combustor Swirl Clocking on Intermediate Turbine Duct Vane Aerodynamics With an Upstream High Pressure Turbine Stage—An Experimental and Computational Study

2016 ◽  
Vol 139 (1) ◽  
Author(s):  
Martin Johansson ◽  
Thomas Povey ◽  
Kam Chana ◽  
Hans Abrahamsson
Keyword(s):  
High Pressure ◽  
Test Facility ◽  
Flow Structures ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Inlet Condition ◽  
Intermediate Turbine Duct ◽  
Turbine Vane ◽  
Flow Through ◽  
Inlet Conditions

Flow in an intermediate turbine duct (ITD) is highly complex, influenced by the upstream turbine stage flow structures, which include tip leakage flow and nonuniformities originating from the upstream high pressure turbine (HPT) vane and rotor. The complexity of the flow structures makes predicting them using numerical methods difficult, hence there exists a need for experimental validation. To evaluate the flow through an intermediate turbine duct including a turning vane, experiments were conducted in the Oxford Turbine Research Facility (OTRF). This is a short duration high speed test facility with a 3/4 engine-sized turbine, operating at the correct nondimensional parameters for aerodynamic and heat transfer measurements. The current configuration consists of a high pressure turbine stage and a downstream duct including a turning vane, for use in a counter-rotating turbine configuration. The facility has the ability to simulate low-NOx combustor swirl at the inlet to the turbine stage. This paper presents experimental aerodynamic results taken with three different turbine stage inlet conditions: a uniform inlet flow and two low-NOx swirl profiles (different clocking positions relative to the high pressure turbine vane). To further explain the flow through the 1.5 stage turbine, results from unsteady computational fluid dynamics (CFD) are included. The effect of varying the high pressure turbine vane inlet condition on the total pressure field through the 1.5 stage turbine, the intermediate turbine duct vane loading, and intermediate turbine duct exit condition are discussed and CFD results are compared with experimental data. The different inlet conditions are found to alter the flow exiting the high pressure turbine rotor. This is seen to have local effects on the intermediate turbine duct vane. With the current stator–stator vane count of 32-24, the effect of relative clocking between the two is found to have a larger effect on the aerodynamics in the intermediate turbine duct than the change in the high pressure turbine stage inlet condition. Given the severity of the low-NOx swirl profiles, this is perhaps surprising.

scholarly journals Commissioning of a Combined Hot-Streak and Swirl Profile Generator in a Transonic Turbine Test Facility

2020 ◽  
Vol 142 (3) ◽  
Author(s):  
Maxwell G. Adams ◽  
Thomas Povey ◽  
Benjamin F. Hall ◽  
David N. Cardwell ◽  
Kam S. Chana ◽  
...  
Keyword(s):  
Boundary Condition ◽  
Reynolds Number ◽  
Gas Turbines ◽  
Test Facility ◽  
Lean Burn ◽  
Numerical Predictions ◽  
Turbine Inlet ◽  
Total Temperature ◽  
Inlet Conditions ◽  
Inlet Boundary Condition

Abstract By enhancing the premixing of fuel and air prior to combustion, recently developed lean-burn combustor systems have led to reduced NOx and particulate emissions in gas turbines. Lean-burn combustor exit flows are typically characterized by nonuniformities in total temperature, or so-called hot-streaks, swirling velocity profiles, and high turbulence intensity. While these systems improve combustor performance, the exiting flow-field presents significant challenges to the aerothermal performance of the downstream turbine. This paper presents the commissioning of a new fully annular lean-burn combustor simulator for use in the Oxford Turbine Research Facility (OTRF), a transonic rotating facility capable of matching nondimensional engine conditions. The combustor simulator can deliver engine-representative turbine inlet conditions featuring swirl and hot-streaks either separately or simultaneously. To the best of our knowledge, this simulator is the first of its kind to be implemented in a rotating turbine test facility.The combustor simulator was experimentally commissioned in two stages. The first stage of commissioning experiments was conducted using a bespoke facility exhausting to atmospheric conditions (Hall and Povey, 2015, “Experimental Study of Non-Reacting Low NOx Combustor Simulator for Scaled Turbine Experiments,” ASME Paper No. GT2015-43530.) and included area surveys of the generated temperature and swirl profiles. The survey data confirmed that the simulator performed as designed, reproducing the key features of a lean-burn combustor. However, due to the hot and cold air mixing process occurring at lower Reynolds number in the facility, there was uncertainty concerning the degree to which the measured temperature profile represented that in OTRF. The second stage of commissioning experiments was conducted with the simulator installed in the OTRF. Measurements of the total temperature field at turbine inlet and of the high-pressure (HP) nozzle guide vane (NGV) loading distributions were obtained and compared to measurements with uniform inlet conditions. The experimental survey results were compared to unsteady numerical predictions of the simulator at both atmospheric and OTRF conditions. A high level of agreement was demonstrated, indicating that the Reynolds number effects associated with the change to OTRF conditions were small. Finally, data from the atmospheric test facility and the OTRF were combined with the numerical predictions to provide an inlet boundary condition for numerical simulation of the test turbine stage. The NGV loading measurements show good agreement with the numerical predictions, providing validation of the stage inlet boundary condition imposed. The successful commissioning of the simulator in the OTRF will enable future experimental studies of lean-burn combustor–turbine interaction.

Flow Field and Hot Streak Migration Through High Pressure Cooled Vanes With Representative Lean Burn Combustor Outflow

2018 ◽  
Author(s):  
T. Bacci ◽  
T. Lenzi ◽  
A. Picchi ◽  
L. Mazzei ◽  
B. Facchini
Keyword(s):  
Experimental Data ◽  
High Pressure ◽  
Flow Field ◽  
Fuel Injection ◽  
Leading Edge ◽  
Flame Stabilization ◽  
Lean Burn ◽  
Aero Engine ◽  
University Of Florence ◽  
Hot Streak

Modern lean burn aero-engine combustors make use of relevant swirl degrees for flame stabilization. Moreover important temperature distortions are generated, in tangential and radial directions, due to discrete fuel injection and liner cooling flows respectively. At the same time, more efficient devices are employed for liner cooling and a less intense mixing with the mainstream occurs. As a result, aggressive swirl fields, high turbulence intensities and strong hot streaks are achieved at the turbine inlet. In order to understand combustor-turbine flow field interactions, it is mandatory to collect reliable experimental data at representative flow conditions. While the separated effects of temperature, swirl and turbulence on the first turbine stage have been widely investigated, reduced experimental data is available when it comes to consider all these factors together. In this perspective, an annular three-sector combustor simulator with fully cooled high pressure vanes has been designed and installed at the THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion cooled liners and six film cooled high pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central NGV aligned with the central swirler. In order to generate representative conditions, a heated mainstream passes though the axial swirlers of the combustor simulator, while the effusion cooled liners are fed by air at ambient temperature. The resulting flow field exiting from the combustor simulator and approaching the cooled vane can be considered representative of a modern Lean Burn aero engine combustor with swirl angles above ±50°, turbulence intensities up to about 28% and maximum-to-minimum temperature ratio of about 1.25. With the final aim of investigating the hot streaks evolution through the cooled high pressure vane, the mean aerothermal field (temperature, pressure and velocity fields) has been evaluated by means of a five hole probe equipped with a thermocouple and traversed upstream and downstream of the NGV cascade.

Evaluation of the Numerical Modelling of a High Pressure Turbine Test Rig

2019 ◽  
Author(s):  
David Luquet ◽  
Francois Julienne ◽  
Aurélien Arntz ◽  
Eric Lippinois
Keyword(s):  
High Pressure ◽  
Infrared Imaging ◽  
Guide Vane ◽  
Fuel Efficiency ◽  
Temporal Variations ◽  
Test Rig ◽  
High Pressure Turbine ◽  
Lean Burn ◽  
Multi Stage ◽  
The Impact

Abstract To improve the fuel efficiency demanded by airlines and regulations, the turbomachinery industry is required to steadily enhance engine performances and numerical prediction capabilities. One of the solutions is the lean burn combustor which dramatically reduces NOx levels compared to rich one. However, one drawback of this technology is its impact on the High-pressure turbine due to large swirl and reduced cooling airflow, inducing large spatial and temporal variations in the turbine inlet condition. This can drastically change the operation of the turbine and our ability to model it using standard practice, usually RANS computation. To investigate this combustor-turbine interaction, the European Commission-funded project FACTOR (Full Aerothermal Combustor-Turbine interactiOns Research) was launched several years ago. A test rig of a combustor simulator coupled with a 1.5 stage turbine was built at a DLR facility. An extensive test campaign comprising 5 holes probes and infrared imaging was performed. These produced an array of aerodynamic quantities at different points of interest along the machine axis. With this project reaching its term by the end of 2017, results have been disseminated to the partners. This allows a comparison of measurements with RANS modeling on this configuration. The present paper deals with this analysis using several RANS computations and the results of the test campaign. First, single row computation of the Nozzle Guide Vane and rotor blade were performed. To impose the boundary conditions, the experimental map were azimuthally averaged to obtain profiles of total temperature, total pressure and flow angles. Second, the impact of some geometrical features was investigated. This was done using the recent addition of unstructured mesh capability in the elsA solver. Finally, multi-stage computations, both steady (mixing plane) and unsteady (sliding mesh) give an insight on the relative accuracy of these interstage models. All these computations were then used to investigate the behavior of this particular turbine. In addition to classical analysis using profiles of averaged data, the loss sources were identified by computing the viscous and thermal entropy production. This paves the way for a better understanding of the possibilities and limitations of our simulation capabilities.

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.

Investigation of High-Pressure Turbine Endwall Film-Cooling Performance Under Realistic Inlet Conditions

2012 ◽  
Vol 28 (4) ◽  
pp. 799-810 ◽  
Author(s):  
Simone Salvadori ◽  
Luca Ottanelli ◽  
Magnus Jonsson ◽  
Peter Ott ◽  
Francesco Martelli
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
High Pressure Turbine ◽  
Cooling Performance ◽  
Inlet Conditions ◽  
Endwall Film Cooling

Influence of Backward- and Forward-Facing Steps on the Flow Through a Turning Mid Turbine Frame

2017 ◽  
Vol 139 (12) ◽  
Author(s):  
Sabine Bauinger ◽  
Emil Goettlich ◽  
Franz Heitmeir ◽  
Franz Malzacher
Keyword(s):  
High Pressure ◽  
Total Pressure ◽  
Test Rig ◽  
Test Setup ◽  
Aero Engine ◽  
Total Temperature ◽  
Flow Through ◽  
Run Up ◽  
Compact Design ◽  
Probe Measurements

For this work, reality effects, more precisely backward-facing steps (BFSs) and forward-facing steps (FFSs), and their influence on the flow through a two-stage two-spool turbine rig under engine-relevant conditions were experimentally investigated. The test rig consists of an high pressure (HP) and an low pressure (LP) stage, with the two rotors rotating in opposite direction with two different rotational speeds. An S-shaped transition duct, which is equipped with turning struts (so-called turning mid turbine frame (TMTF)) and making therefore a LP stator redundant, connects both stages and leads the flow from a smaller to a larger diameter. This test setup allows the investigation of a TMTF deformation, which occurs in a real aero-engine due to non-uniform warming of the duct during operation—especially during run up—and causes BFSs and FFSs in the flow path. This happens for nonsegmented ducts, which are predominantly part of smaller engines. In the case of the test rig, steps were not generated by varying temperature but by shifting the TMTF in horizontal direction while the rotor and its casing were kept in the same position. In this way, both BFSs and FFSs between duct endwalls and rotor casing could be created. In order to avoid steps further downstream of the interface between HP rotor and TMTF, the complete aft rig was moved laterally too. In this case, the aft rig incorporates among others the LP rotor, the LP rotor casing, and the deswirler downstream of the LP stage. In order to catch the influence of the steps on the whole flow field, 360 deg rake traverses were performed downstream of the HP rotor, downstream of the duct, and downstream of the LP rotor with newly designed, laser-sintered combi-rakes for the measurement of total pressure and total temperature. Only the compact design of the rakes, which can be easily realized by additive manufacturing, makes the aforementioned 360 deg traverses in this test rig possible and allows a number of radial measurements positions, which is comparable to those of a five-hole probe. To get a more detailed information about the flow, also five-hole probe measurements were carried out in three measurement planes and compared to the results of the combi-rakes.

Experimental Investigation of Vane Clocking in a One and 1/2 Stage High Pressure Turbine

2004 ◽  
Author(s):  
Charles W. Haldeman ◽  
Michael G. Dunn ◽  
John W. Barter ◽  
Brian R. Green ◽  
Robert F. Bergholz
Keyword(s):  
High Pressure ◽  
Low Pressure ◽  
Test Facility ◽  
Maximum Efficiency ◽  
High Pressure Turbine ◽  
Data Set ◽  
Time Resolved ◽  
Low Pressure Turbine ◽  
Turbine Vane ◽  
Vane Clocking

Aerodynamic measurements were acquired on a modern single-stage, transonic, high-pressure turbine with the adjacent low-pressure turbine vane row (a typical civilian one and one-half stage turbine rig) to observe the effects of low-pressure turbine vane clocking on overall turbine performance. The turbine rig (loosely referred to in this paper as the stage) was operated at design corrected conditions using the Ohio State University Gas Turbine Laboratory Turbine Test Facility (TTF). The research program utilized uncooled hardware in which all three airfoils were heavily instrumented at multiple spans to develop a full clocking dataset. The low-pressure turbine vane row (LPTV) was clocked relative to the high-pressure turbine vane row (HPTV). Various methods were used to evaluate the influence of clocking on the aeroperformance (efficiency) and the aerodynamics (pressure loading) of the LPTV, including time-resolved and time-averaged measurements. A change in overall efficiency of approximately 2–3% due to clocking effects is demonstrated and could be observed using a variety of independent methods. Maximum efficiency is obtained when the time-average surface pressures are highest on the LPTV and the time-resolved surface pressure (both in the time domain and frequency domain) show the least amount of variation. The overall effect is obtained by integrating over the entire airfoil, as the three-dimensional effects on the LPTV surface are significant. This experimental data set validates several computational research efforts that suggested wake migration is the primary reason for the perceived effectiveness of vane clocking. The suggestion that wake migration is the dominate mechanism in generating the clocking effect is also consistent with anecdotal evidence that fully cooled engine rigs do not see a great deal of clocking effect. This is consistent since the additional disturbances induced by the cooling flows and/or the combustor make it extremely difficult to find an alignment for the LPTV given the strong 3D nature of modern high-pressure turbine flows.

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