Aerodynamic design of a wave rotor to high pressure turbine transition duct

1998 ◽  
Author(s):  
Steven Gegg
Keyword(s):  
High Pressure ◽  
Aerodynamic Design ◽  
Wave Rotor ◽  
High Pressure Turbine

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.

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.

Validation of MISES 2-D Boundary Layer Code for High Pressure Turbine Aerodynamic Design

2007 ◽  
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 ◽  
Linear Cascade ◽  
Reduce Operating Cost ◽  
Wide Range

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 which 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 2-D linear cascade approximating the performance of a moderately-loaded, mid-pitch section from a modern aircraft high-pressure turbine [1–2]. 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° to +14.5° relative to design incidence.

A dynamic probabilistic design method for blade-tip radial running clearance of aeroengine high-pressure turbine

2014 ◽  
Vol 229 (10) ◽  
pp. 1861-1872 ◽  
Author(s):  
Cheng-Wei Fei ◽  
Wen-Zhong Tang ◽  
Guang-chen Bai ◽  
Zhi-Ying Chen
Keyword(s):  
High Pressure ◽  
Response Surface ◽  
Response Surface Method ◽  
Probabilistic Analysis ◽  
High Reliability ◽  
Probabilistic Design ◽  
High Pressure Turbine ◽  
Analysis And Design ◽  
Surface Method ◽  
Blade Tip

Around the engineering background of the probabilistic design of high-pressure turbine (HPT) blade-tip radial running clearance (BTRRC) which conduces to the high-performance and high-reliability of aeroengine, a distributed collaborative extremum response surface method (DCERSM) was proposed for the dynamic probabilistic analysis of turbomachinery. On the basis of investigating extremum response surface method (ERSM), the mathematical model of DCERSM was established. The DCERSM was applied to the dynamic probabilistic analysis of BTRRC. The results show that the blade-tip radial static clearance δ = 1.82 mm is advisable synthetically considering the reliability and efficiency of gas turbine. As revealed by the comparison of three methods (DCERSM, ERSM, and Monte Carlo method), the DCERSM reshapes the possibility of the probabilistic analysis for turbomachinery and improves the computational efficiency while preserving computational accuracy. The DCERSM offers a useful insight for BTRRC dynamic probabilistic analysis and optimization. The present study enrichs mechanical reliability analysis and design theory.

Influence of Hot Streak Temperature Ratio on Low Pressure Stage of a Vaneless Counter-Rotating Turbine

2007 ◽  
Author(s):  
Qingjun Zhao ◽  
Fei Tang ◽  
Huishe Wang ◽  
Jianyi Du ◽  
Xiaolu Zhao ◽  
...  
Keyword(s):  
Shock Wave ◽  
High Pressure ◽  
Secondary Flow ◽  
Low Pressure ◽  
Turbine Rotor ◽  
Temperature Ratio ◽  
High Pressure Turbine ◽  
Pressure Stage ◽  
Low Pressure Turbine ◽  
Hot Streak

In order to explore the influence of hot streak temperature ratio on low pressure stage of a Vaneless Counter-Rotating Turbine, three-dimensional multiblade row unsteady Navier-Stokes simulations have been performed. The predicted results show that hot streaks are not mixed out by the time they reach the exit of the high pressure turbine rotor. The separation of colder and hotter fluids is observed at the inlet of the low pressure turbine rotor. After making interactions with the inner-extending shock wave and outer-extending shock wave in the high pressure turbine rotor, the hotter fluid migrates towards the pressure surface of the low pressure turbine rotor, and the most of colder fluid migrates to the suction surface of the low pressure turbine rotor. The migrating characteristics of the hot streaks are predominated by the secondary flow in the low pressure turbine rotor. The effect of buoyancy on the hotter fluid is very weak in the low pressure turbine rotor. The results also indicate that the secondary flow intensifies in the low pressure turbine rotor when the hot streak temperature ratio is increased. The effects of the hot streak temperature ratio on the relative Mach number and the relative flow angle at the inlet of the low pressure turbine rotor are very remarkable. The isentropic efficiency of the Vaneless Counter-Rotating Turbine decreases as the hot streak temperature ratio is increased.

Large-Scale Simulation of the Clocking Impact of 2D Combustor Profile on a Two Stage High Pressure Turbine

2014 ◽  
Author(s):  
Thomas E. Dyson ◽  
David B. Helmer ◽  
James A. Tallman
Keyword(s):  
High Pressure ◽  
Large Scale ◽  
High Pressure Turbine ◽  
Cfd Simulations ◽  
Two Stage ◽  
Substantial Impact ◽  
Sliding Mesh ◽  
Wall Flow ◽  
End Wall ◽  
Film Flows

This paper presents sliding-mesh unsteady CFD simulations of high-pressure turbine sections of a modern aviation engine in an extension of previously presented work [1]. The simulation included both the first and second stages of a two-stage high-pressure turbine. Half-wheel domains were used, with source terms representing purge and film flows. The end-wall flow-path cavities were incorporated in the domain to a limited extent. The passage-to-passage variation in thermal predictions was compared for a 1D and 2D turbine inlet boundary condition. Substantial impact was observed on both first and second stage vanes despite the mixing from the first stage blade. Qualitative and quantitative differences in surface temperature distributions were observed due to different ratios between airfoil counts in the two domains.

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.

Aerodynamic Efficiency Optimization of the 1st Stage of Transonic High Pressure Turbine through Lean and Sweep Angles

2019 ◽  
Vol 36 (3) ◽  
pp. 245-256
Author(s):  
Yoonki Kim ◽  
Sanga Lee ◽  
Kwanjung Yee ◽  
Young-Seok Kang
Keyword(s):  
High Pressure ◽  
Surrogate Model ◽  
Design Space ◽  
Latin Hypercube Sampling ◽  
Expected Improvement ◽  
Total Efficiency ◽  
High Pressure Turbine ◽  
Sweep Angle ◽  
Design Variables ◽  
Optimal Latin Hypercube Sampling

Abstract The purpose of this study is to optimize the 1st stage of the transonic high pressure turbine (HPT) for enhancement of aerodynamic performance. Isentropic total-to-total efficiency is designated as the objective function. Since the isentropic efficiency can be improved through modifying the geometry of vane and rotor blade, lean angle and sweep angle are chosen as design variables, which can effectively alter the blade geometry. The sensitivities of each design variable are investigated by applying lean and sweep angles to the base nozzle and rotor, respectively. The design space is also determined based on the results of the parametric study. For the design of experiment (DoE), Optimal Latin Hypercube sampling is adopted, so that 25 evenly distributed samples are selected on the design space. Sequentially, based on the values from the CFD calculation, Kriging surrogate model is constructed and refined using Expected Improvement (EI). With the converged surrogate model, optimum solution is sought by using the Genetic Algorithm. As a result, the efficiency of optimum turbine 1st stage is increased by 1.07 % point compared to that of the base turbine 1st stage. Also, the blade loading, pressure distribution, static entropy, shock structure, and secondary flow are thoroughly discussed.

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