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.

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.

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.

Validation of MISES Two-Dimensional Boundary Layer Code for High-Pressure Turbine Aerodynamic Design

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
Philip L. Andrew ◽  
Harika S. Kahveci
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Operating Cost ◽  
Design Tool ◽  
Operating Conditions ◽  
Aerodynamic Design ◽  
High Pressure Turbine ◽  
Reduce Operating Cost ◽  
Wide Range ◽  
Dimensional Boundary

Avoiding aerodynamic separation and excessive shock losses in gas turbine turbomachinery components can reduce fuel usage and thus reduce operating cost. In order to achieve this, blading designs should be made robust to a wide range of operating conditions. Consequently, a design tool is needed—one that can be executed quickly for each of many operating conditions and on each of several design sections, which will accurately capture loss, turning, and loading. This paper presents the validation of a boundary layer code, MISES, versus experimental data from a 2D linear cascade approximating the performance of a moderately loaded mid-pitch section from a modern aircraft high-pressure turbine. The validation versus measured loading, turning, and total pressure loss is presented for a range of exit Mach numbers from ≈0.5 to 1.2 and across a range of incidence from −10 deg to +14.5 deg relative to design incidence.

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

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.

scholarly journals The Design and Evaluation of a High Pressure Ratio Radial Turbine

1992 ◽  
Author(s):  
K. R. Pullen ◽  
N. C. Baines ◽  
S. H. Hill
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Performance Measurements ◽  
Turbine Engine ◽  
High Pressure Ratio ◽  
Turbine Efficiency ◽  
Wide Range ◽  
Dimensional Parameters ◽  
Very High

A single stage, high speed, high pressure ratio radial inflow turbine was designed for a single shaft gas turbine engine in the 200 kW power range. A model turbine has been tested in a cold rig facility with correct simulation of the important non-dimensional parameters. Performance measurements over a wide range of operation were made, together with extensive volute and exhaust traverses, so that gas velocities and incidence and deviation angles could be deduced. The turbine efficiency was lower than expected at all but the lowest speed. The rotor incidence and exit swirl angles, as obtained from the rig test data, were very similar to the design assumptions. However, evidence was found of a region of separation in the nozzle vane passages, presumably caused by a very high curvature in the endwall just upstream of the vane leading edges. The effects of such a separation are shown to be consistent with the observed performance.

Aero-thermodynamic and chemical process interactions in an axial high-pressure turbine of aircraft engines

2018 ◽  
Vol 20 (6) ◽  
pp. 653-669 ◽  
Author(s):  
Trung Hieu Nguyen ◽  
Phuong Nguyen-Tri ◽  
Xavier Vancassel ◽  
Francois Garnier
Keyword(s):  
High Pressure ◽  
Three Dimensional ◽  
Chemical Processes ◽  
Rotor Blades ◽  
Rotor Speed ◽  
High Pressure Turbine ◽  
Three Dimensional Modeling ◽  
Dimensional Modeling ◽  
Speed Effect ◽  
Flow Vortex

Precise investigation of aero-thermodynamic and chemical processes relating to environmental precursor pollutants in an aircraft turbine is challenging because of the complexity of transformation processes at high temperature and high pressure. We present here, for the first time, new insights into the study of aero-thermodynamic processes, formation of nitrate and sulfate aerosol precursors, and investigate the influence of chemical processes on aero-thermodynamics. We also shed light on the effect of three-dimensional blade profile, radial spacing, and rotor speed on the performance of a high-pressure turbine. We highlight that flow vortex and the variation of chemical formation which appear in both rear stator blades and rear rotor blades. We found that the chemical processes were affected by the evolution of temperature (maximum of 16.9%) and flow velocity (maximum of 38.8%). Contrary to the conservative one-dimensional and two-dimensional modeling, which provide only the flow trends and flow evolution at cylindrical surface, respectively, our three-dimensional modeling approach offers the possibility of combining information on radial spacing and rotor-speed effect by providing three-dimensional images of spatial-geometry effect on aero-thermodynamic and chemical processes. Quantitatively, the magnitude of change in aero-thermodynamics and nitrogen oxidation may be expected to be up to 17% and 48%, respectively, over a stage of the high-pressure turbine.

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.

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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