Integration of a transonic high-pressure turbine with a rotating detonation combustor and a diffuser

2020 ◽  
Vol 0 (0) ◽  
Author(s):  
Zhe Liu ◽  
James Braun ◽  
Guillermo Paniagua
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
High Efficiency ◽  
Navier Stokes ◽  
High Pressure Turbine ◽  
Turbine Efficiency ◽  
Main Challenge ◽  
Reynolds Averaged Navier Stokes ◽  
Inlet Conditions ◽  
Rotating Detonation

AbstractIn this paper, a diffuser is used to integrate a transonic high-pressure turbine with a rotating detonation combustor (RDC). The paper focuses on the required design modifications to the turbine endwalls (EW) to enable high efficiency, while preserving the airfoil blade-to-blade geometry. The main challenge is the stator passage unstarting, due to the high inlet Mach number. First of all, steady Reynolds Averaged Navier Stokes simulations were performed to compare the efficiency of turbines with constant-radius EWs to turbines with axisymmetric EWs. A modified EW design prevented the unstarting of the stator passage, enabling a significant gain in performance. Afterward, the influence on the turbine efficiency and damping due to the unsteadiness from the diffuser-like fluctuations of the RDC was evaluated with unsteady Reynolds Averaged Navier Stokes simulations with a mixing plane approach (MPA). Full unsteady simulations were carried out on selected inlet conditions and compared to the mixing plane results. This parametric study provides turbine designers with recommended diffusion rates along the vane EWs. Additionally, we provide guidance on the upstream diffuser design, specifically the required damping and outlet Mach number.

InterTurb: High-Pressure Turbine Rig for the Investigation of Combustor-Turbine Interaction

2017 ◽  
Author(s):  
T. Wolf ◽  
K. Lehmann ◽  
L. Willer ◽  
A. Pahs ◽  
M. Rößling ◽  
...  
Keyword(s):  
High Pressure ◽  
Rotor Blades ◽  
Aerodynamic Design ◽  
High Pressure Turbine ◽  
Turbine Efficiency ◽  
Testing Facility ◽  
Dimensional Parameters ◽  
Inlet Conditions ◽  
Secondary Air

This paper introduces a new 2-stage high-pressure turbine rig for aerodynamic investigations. It is operated by DLR Göttingen (Germany) and installed in DLR’s new testing facility NG-Turb. The rig’s geometrical size as well as the non-dimensional parameters are comparable to a modern engine in the small to medium thrust range. The turbine rig closely resembles engine hardware and features all relevant blade and vane cooling as well as secondary air-system flows. The effect of variations of each individual flow and different tip clearances on overall turbine efficiency will be studied. While the first part of the testing program will be based on uniform inlet conditions the second part will be run with a combustor simulator, which is based on electrical heaters and delivers a flow field similar to a rich-burn combustor. In order to find the optimum relative position for maximum turbine efficiency the combustor simulator can be rotated relative to the HPT inlet (clocking). For the same reasons the stators can also be clocked. The paper gives a brief overview of the testing facility and from there on focuses on the HPT rig features such as aerodynamic design, cooling and sealing flows. The aerodynamic optimisation of the stator vanes and shroudless rotor blades will be outlined. Further topics are the aerodynamic design of the combustor simulator, a comparison with engine combustors as well as the implementation in the rig. The paper also describes the rig instrumentation in the stationary and rotating system which most importantly focuses on measurements of efficiency and capturing of traverse data. The topic of blade and vane manufacturing via direct metal laser sintering will be briefly covered. The discussion of test results and comparison with numerical simulations will be the subject of a follow-up paper.

scholarly journals High-Fidelity Simulations of a High-Pressure Turbine Vane Subject to Large Disturbances: Effect of Exit Mach Number on Losses

2021 ◽  
pp. 1-11
Author(s):  
Yaomin Zhao ◽  
Richard Sandberg
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Mach Number ◽  
Strong Shock ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Loss Analysis ◽  
Turbine Efficiency ◽  
Turbulence Condition ◽  
Exit Mach Number

Abstract We report on a series of highly resolved large-eddy simulations of the LS89 high-pressure turbine (HPT) vane, varying the exit Mach number between Ma=0.7 and 1.1. In order to accurately resolve the blade boundary layers and enforce pitchwise periodicity, we for the first time use an overset mesh method, which consists of an O-type grid around the blade overlapping with a background H-type grid. The simulations were conducted either with a synthetic inlet turbulence condition or including upstream bars. A quantitative comparison shows that the computationally more efficient synthetic method is able to reproduce the turbulence characterictics of the upstream bars. We further perform a detailed analysis of the flow fields, showing that the varying exit Mach number significantly changes the turbine efficiency by affecting the suction-side transition, blade boundary layer profiles, and wake mixing. In particular, the Ma=1.1 case includes a strong shock that interacts with the trailing edge, causing an increased complexity of the flow field. We use our recently developed entropy loss analysis (Zhao and Sandberg, GT2019-90126) to decompose the overall loss into different source terms and identify the regions that dominate the loss generation. Comparing the different Ma cases, we conclude that the main mechanism for the extra loss generation in the Ma=1.1 case is the shock-related strong pressure gradient interacting with the turbulent boundary layer and the wake, resulting in significant turbulence production and extensive viscous dissipation.

scholarly journals High-Fidelity Simulations of a High-Pressure Turbine Vane Subject to Large Disturbances: Effect of Exit Mach Number on Losses

2020 ◽  
Author(s):  
Yaomin Zhao ◽  
Richard D. Sandberg
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Mach Number ◽  
Strong Shock ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Loss Analysis ◽  
Turbine Efficiency ◽  
Turbulence Condition ◽  
Exit Mach Number

Abstract We report on a series of highly resolved large-eddy simulations of the LS89 high-pressure turbine (HPT) vane, varying the exit Mach number between Ma = 0.7 and 1.1. In order to accurately resolve the blade boundary layers and enforce pitchwise periodicity, we for the first time use an overset mesh method, which consists of an O-type grid around the blade overlapping with a background H-type grid. The simulations were conducted either with a synthetic inlet turbulence condition or including upstream bars. A quantitative comparison shows that the computationally more efficient synthetic method is able to reproduce the turbulence characterictics of the upstream bars. We further perform a detailed analysis of the flow fields, showing that the varying exit Mach number significantly changes the turbine efficiency by affecting the suction-side transition, blade boundary layer profiles, and wake mixing. In particular, the Ma = 1.1 case includes a strong shock that interacts with the trailing edge, causing an increased complexity of the flow field. We use our recently developed entropy loss analysis (Zhao and Sandberg, GT2019-90126) to decompose the overall loss into different source terms and identify the regions that dominate the loss generation. Comparing the different Ma cases, we conclude that the main mechanism for the extra loss generation in the Ma = 1.1 case is the shock-related strong pressure gradient interacting with the turbulent boundary layer and the wake, resulting in significant turbulence production and extensive viscous dissipation.

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

Design of an Annular-Radial Diffuser for Operation With a Supercritical CO2 Radial Inflow Turbine

2019 ◽  
Vol 141 (8) ◽  
Author(s):  
Joshua A. Keep ◽  
Ingo H. J. Jahn
Keyword(s):  
Supercritical Co2 ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Geometric Feature ◽  
Small Scale ◽  
Rear Side ◽  
Power Cycles ◽  
Reynolds Averaged Navier Stokes ◽  
Inlet Conditions ◽  
Conventional Manner

Radial inflow turbines are a relevant architecture for energy extraction from supercritical CO2 power cycles for scales less than 10 MW. To ensure stage and overall cycle efficiency, it is desirable to recover exhaust energy from the turbine stage through the inclusion of a suitable diffuser in the turbine exhaust stream. In supercritical CO2 Brayton cycles, the high turbine inlet pressure can lead to sealing challenges at small scale if the rotor is supported from the rotor rear side in the conventional manner. An alternative is a layout where the rotor exit faces the bearing system. While such a layout is attractive for the sealing system, it limits the axial space claim of the diffuser. Designs of a combined annular-radial diffuser are considered as a means to meet the aforementioned packaging challenges of this rotor layout. Diffuser performance is assessed numerically with the use of Reynolds-averaged Navier--Stokes (RANS) and unsteady Reynolds-averaged Navier--Stokes (URANS) calculations. To appropriately account for cross coupling with the stage, a single blade passage of the entire stage is modeled. Assessment of diffuser inlet conditions, and off-design performance analysis, reveals that the investigated diffuser designs are performance robust to high swirl, high inlet blockage, and highly nonuniform mass flux distribution. Diffuser component performance is dominated by the annular-radial bend. The incorporation of a constant sectional area bend is the key geometric feature in rendering the highly nonuniform turbine exit flow (dominated by tip clearance flows at the shroud) more uniform.

Using CFD to Reduce Resonant Stresses on a Single-Stage, High-Pressure Turbine Blade

2002 ◽  
Author(s):  
J. P. Clark ◽  
A. S. Aggarwala ◽  
M. A. Velonis ◽  
R. E. Gacek ◽  
S. S. Magge ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Guide Vane ◽  
Turbine Blades ◽  
Suction Side ◽  
Lessons Learned ◽  
Single Stage ◽  
Navier Stokes ◽  
High Pressure Turbine ◽  
Time Resolved

The ability to predict levels of unsteady forcing on high-pressure turbine blades is critical to avoid high-cycle fatigue failures. In this study, 3D time-resolved computational fluid dynamics is used within the design cycle to predict accurately the levels of unsteady forcing on a single-stage high-pressure turbine blade. Further, nozzle-guide-vane geometry changes including asymmetric circumferential spacing and suction-side modification are considered and rigorously analyzed to reduce levels of unsteady blade forcing. The latter is ultimately implemented in a development engine, and it is shown successfully to reduce resonant stresses on the blade. This investigation builds upon data that was recently obtained in a full-scale, transonic turbine rig to validate a Reynolds-Averaged Navier-Stokes (RANS) flow solver for the prediction of both the magnitude and phase of unsteady forcing in a single-stage HPT and the lessons learned in that study.

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.

scholarly journals Effect of Inlet Temperature Non-Uniformity on High-Pressure Turbine Performance

2010 ◽  
Author(s):  
Craig I. Smith ◽  
Dongil Chang ◽  
Stavros Tavoularis
Keyword(s):  
High Pressure ◽  
Heat Load ◽  
Stokes Equations ◽  
Computation Time ◽  
Radial Pressure ◽  
Navier Stokes ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Navier Stokes Equations ◽  
Hot Streak

The temperature of the flow entering a high-pressure turbine stage is inherently non-uniform, as it is produced by several discrete, azimuthally-distributed combustors. In general, however, industrial simulations assume inlet temperature uniformity to simplify the preparation process and reduce computation time. The effects of a non-uniform inlet field on the performance of a commercial, transonic, single-stage, high-pressure, axial turbine with a curved inlet duct have been investigated numerically by performing URANS (Unsteady Reynolds-Averaged Navier-Stokes equations) simulations with the SST (Shear Stress Transport) turbulence model. By adjusting the alignment of the experimentally-based inlet temperature field with respect to the stator vanes, two clocking configurations were generated: an aligned case, in which each hot streak impinged on a vane and a misaligned case, in which each hot streak passed between two vanes. In the aligned configuration, the hot streaks produced higher time-averaged heat load on the vanes and lower heat load on the blades. As the aligned hot streaks impinged on the stator vanes, they also spread spanwise due to the actions of the casing passage vortices and the radial pressure gradient; this resulted in a stream entering the rotor with relatively low temperature variations. The misaligned hot streaks were convected undisturbed past the relatively cool vane section. Relatively high time-averaged enthalpy values were found to occur on the pressure side of the blades in the misaligned configuration. The non-uniformity of the time-averaged enthalpy on the blade surfaces was lower in the aligned configuration. The flow exiting the rotor section was much less non-uniform in the aligned case, but differences in calculated efficiency were not significant.

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.

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