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.

Multi-Disciplinary Analyses for the Design of a High Pressure Turbine Blade Tip

2016 ◽  
Author(s):  
Stefano Caloni ◽  
Shahrokh Shahpar
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Cooling Flow ◽  
Barrier Coating ◽  
Flow Features ◽  
Internally Cooled ◽  
Blade Tip

The design of a high pressure turbine blade is a challenging task requiring multiple disciplines to be solved simultaneously. Most recently, conjugate analyses are being developed to tackle such a problem; they are able to resolve both the fluid dynamics in a turbine passage and the thermal distribution in the solid part of the component. In this paper, the in-house Hydra CFD solver is used to analyse a high pressure shroudless turbine blade for a modern jet engine. The turbine is internally cooled and a Thermal Barrier Coating (TBC) is applied on the aerofoil surface. The coupling technique used at the interface in the presence of the TBC is described. The flow features at the tip of the turbine blade are the main focus of this study. Four different tip configurations are analysed. A flat tip and a squealer tip are chosen as reference designs; however the effects of opening the Trailing Edge (TE) on the Suction Side (SS) and the Pressure Side (PS) are also investigated. Both a cooled and an uncooled configuration of the turbine blade are analysed and the effect of the cooling flow on the over tip leakage is studied. Finally, conjugate analyses for the cooled turbine blades are used to predict the temperature reached by the different tip designs. The design with an opened TE on the SS shows a significant aerodynamic improvement over the others without increasing the temperature the tip has to withstand in operation.

Time-Averaged and Time-Accurate Aero-Dynamics for the Recessed Tip Cavity of a High-Pressure Turbine Blade and the Outer Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Turbine Rotor ◽  
Single Stage ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a 3D, non-linear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

Numerical investigation of high-pressure turbine blade tip flows: analysis of aerodynamics

2011 ◽  
Vol 225 (7) ◽  
pp. 940-953 ◽  
Author(s):  
P Vass ◽  
T Arts
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Suction Side ◽  
Point Of View ◽  
Internal Cooling ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Design Solutions ◽  
Exit Mach Number

The current contribution reports on the validation and analysis of three-dimensional computational results of the flow around four distinct high-pressure turbine blade tip geometries (TG1, 2, 3, and 4 hereinafter), taking into account the effect of the entire internal cooling setup inside the blade, at design exit Mach number: M = 0.8, and high exit Reynolds number: Re C = 900 000. Three of the four geometries represent different tip design solutions – TG1: full squealer rim; TG2: single squealer on the suction side; TG3: partial suction and pressure side squealer, and one (TG4) models TG1 in worn condition. This article provides a comparison between the different geometries from the aerodynamic point of view, analyses the losses, and evaluates the distinct design solutions. An assessment of the effect of the uneven rubbing of the blade tip was performed as well. TG1 was found to be the top performer followed by TG3 and TG2. According to the investigations, the effect of rubbing increased the kinetic loss coefficient by 10–15 per cent.

An Investigation on Aerodynamics Loss Mechanism of Squealer Tips of a High Pressure Turbine Blade Using URANS

2016 ◽  
Author(s):  
Jin-sol Jung ◽  
Okey Kwon ◽  
Changmin Son
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Wall Thickness ◽  
Suction Side ◽  
Leakage Flow ◽  
Pressure Side ◽  
Tip Leakage Flow ◽  
High Pressure Turbine ◽  
Loss Mechanism ◽  
Tip Leakage

The flow leaking over the tip of a high pressure turbine blade generates significant aerodynamic losses as it mixes with the mainstream flow. This study investigates the effect of blade tip geometries on turbine performance with both steady RANS and unsteady URANS analyses. Five different squealer geometries for a high pressure turbine blade have been examined: squealer on pressure side, squealer on suction side, cavity squealer, cavity squealer with pressure side cutback, and cavity squealer with suction side cutback. With the case of the cavity squealer, three different squealer wall thickness are investigated for the wall thickness (w) of 1x, 2x and 4x of the tip gap (G). The unsteady flow analyses using CFX have been conducted to investigate unsteady characteristics of the tip leakage flow and its influence on turbine performances. Through the comparison between URANS analyses, detailed vortex and wake structures are identified and studied at different fidelities. It is found that the over tip leakage flow loss is affected by the tip suction side geometry rather than that of the pressure side geometry. The unsteady results have contributed to resolve the fundamentals of vortex structures and aerodynamic loss mechanisms in a high pressure turbine stage.

Averaged and Time-Dependent Aerodynamics of a High Pressure Turbine Blade Tip Cavity and Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Vol 127 (4) ◽  
pp. 736-746 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Single Stage ◽  
Time Dependent ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high-frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a three-dimensional (3D), nonlinear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

Gradient-Based Adjoint and Design of Experiment CFD Methodologies to Improve the Manufacturability of High Pressure Turbine Blades

2016 ◽  
Author(s):  
Giulio Zamboni ◽  
Gabriel Banks ◽  
Simon Bather
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Aerodynamic Performance ◽  
Correlation Factor ◽  
Control Points ◽  
High Pressure Turbine ◽  
Flow Capacity ◽  
Profile Tolerance ◽  
The Impact

The tolerance of a turbine blade aerofoil is determined by the requirements to achieve an aerodynamic performance in operation. In fact, the manufacturing tolerance applied to the profile is driven by the effects of geometrical non-conformances on the efficiency and flow capacity of the aerofoil. However, this tolerance also has an impact on the ease with which the aerofoil can be manufactured, with tighter tolerance leading to lower manufacturing conformity. This paper details the application of an adjoint RANS solver and the according series of Design of Experiments (DoE) CFD calculations for a high pressure turbine blade to the above problem. There are two aims of this work; the first is to show that simpler linear CFD perturbation can be used to evaluate the effect of the geometric non-conformance. The second is to validate the spatial geometric correlation factor of the control points used in the manufacturing process on the performance evaluation with DoE techniques. This also verified the applicability of the adjoint CFD techniques; in fact the adjoint CFD calculation is an order of magnitude less computationally expensive than a large series of DoE RANS CFD calculations. The results confirm that the peak suction area is the most critical control region for the effect on the efficiency and flow capacity. Moreover, the CFD investigations show that a significant level of correlation exists between the influence factors at different control points. This suggests that not only the amount of geometric deviation but also the stream surface variation of profile tolerance significantly influence the final aerodynamic performance. The results from this calculation allow the creation of a 3D sensitivity map which will be used during the manufacturing of the aerofoil to optimise the control of the spatial distribution of the geometric non-conformance and to directly assess the expected performance effect during the manufacturing quality inspection. The methodology detailed in this paper shows how the CFD adjoint methods could be used for improved manufacturability of turbine blades ensuring that the critical characteristic features are controlled on the surface, relaxing the profile tolerance on those surface areas where the impact on the aerodynamic performance is predicted to be lower.

scholarly journals Computational Modeling and Thermal Paint Verification of Film-Cooling Designs for an Unshrouded High-Pressure Turbine Blade

1999 ◽  
Author(s):  
Steven G. Gegg ◽  
Nathan J. Heidegger ◽  
Ronald A. Mikkelson
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
Data Exchange ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Extreme High Temperature ◽  
Blade Tip

High pressure turbine blades are exposed to an extreme high temperature environment due to increasing turbine inlet temperature. High heat fluxes are likely on the blade pressure surface. Other regions, such as the trailing edge and blade tip may be difficult to cool uniformly. Unshrouded blades present an additional challenge due to the pressure driven transport of hot gas across the blade tip. The blade tip region is therefore prone to severe thermal stress, fatigue and oxidation. In order to develop effective cooling methods, designers require detailed flow and heat transfer information. This paper reports on computational aerodynamics and heat transfer studies for an unshrouded high pressure turbine blade. The emphasis is placed on the application of appropriate 3-D models for the prediction of airfoil surface temperatures. Details of the film cooling model, boundary conditions and data exchange with heat transfer models are described. The analysis approach has been refined for design use to provide timely and accurate results. Film cooling designs are to be tailored for the best coverage of the blade tip region. Designs include near-tip pressure side films and blade tip cooling holes. Hole placement and angle are investigated to achieve the best coolant coverage on the blade tip. Analytical results are compared to a thermal paint test on engine hardware. In addition to film cooling strategies, other aerodynamic/heat transfer design approaches are discussed to address the cooling requirements for an unshrouded blade.

Aerodynamic Performance Investigation of Turbine in the Event of One Blade Primary Fracture Failure

2020 ◽  
Author(s):  
Shenghui Zhang ◽  
Shuiting Ding ◽  
Tian Qiu
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
High Pressure Turbine ◽  
3 Dimensional ◽  
Fracture Failure ◽  
Pressure Distributions ◽  
Mach Number Distribution ◽  
Blade Failure

Abstract One of major safety requirements from current airworthiness regulations is that the probability of hazardous engine effects should not occur exceed 10−7 per engine flight hour even in the event of component failure. Service experience of aeroengines indicates that turbine blade fracture is a common fault whose probability is far more than 10−7 per engine flight hour. It is obvious that overall engine system will be affected by blade failure. So, aerodynamic performance investigation in the event of one blade fracture failure has been assessed in the current study. With ANSYS-CFX, numerical model of GE-E3 (Energy Efficient Engine) high pressure turbine was established according to literature data. By comparing surface Mach number distribution at mid-span of vane in the first stage obtained numerically and experimentally, the most efficient turbulence model, i.e., the SST k-ω model, was identified. Based on the model, the 3-dimensional flow simulations under two configurations, full wheel geometry GE-E3 high pressure turbine without and with one blade fracture failure have been achieved. The following conclusions were drawn from 3-dimensional simulations: firstly, as for GE-E3 high pressure turbine, the effect of single turbine blade failure on turbine characteristics is slight; secondly, with blade loading coefficient as a criterion which is used for judging whether blade is affected, five blades which are significantly affected can be identified, and the surface pressure distributions of these five affected blades alter to varying degrees, accord-ingly, these film outflow static pressure characteristics alter as well; thirdly, after turbine blade fails, airflow accelerates violently along the suction side of downstream blade closest to failed blade and separates, however, air flow can not expand efficiently along the pressure side of upstream blade nearest to failed blade.

10 Years Later: A Technical and Financial Review of the United States Navy’s High Pressure Turbine Blade Refurbishment Program

2010 ◽  
Author(s):  
Rob Neff ◽  
Matthew Driscoll
Keyword(s):  
United States ◽  
High Pressure ◽  
Turbine Blade ◽  
Financial Analysis ◽  
Risk Mitigation ◽  
Turbine Blades ◽  
The United States ◽  
High Pressure Turbine ◽  
Financial Assessment ◽  
The Cost

In 1999, the United States Navy implemented an LM2500 High Pressure Turbine Blade Refurbishment Program. Traditionally, the US Navy had replaced high pressure turbine components each time an engine was removed from a ship during its depot overhaul visit. Following successful testing of several Rainbow rotors built up with refurbished LM2500 blades, as well as experience gained by the Royal Australian Navy, refurbishment of stage 1 and 2 high pressure turbine blades was adopted vice the replace with new part strategy previously utilized. This paper takes a fresh look at the blade refurbishment effort from two perspectives, first, an updated technical assessment is made of Rainbow rotor components as well as parts which were implemented as part of the refurbishment program to evaluate their current (2009) condition and define service life expectations. Secondly, a financial assessment is made of the program itself, defining the cost avoidance of refurbishing customer owned blades versus the cost to procure new components. The financial analysis will also include commentary on risk mitigation based upon the hundreds of thousands of operating hours on these components have acquired while deployed at sea.

Robust Design Optimization Applied to a High Pressure Turbine Blade Based on Surrogate Modelling Techniques

2015 ◽  
Author(s):  
Frank Wagner ◽  
Arnold Kühhorn ◽  
Roland Parchem
Keyword(s):  
High Pressure ◽  
Design Optimization ◽  
Turbine Blade ◽  
Robust Design ◽  
Turbine Blades ◽  
Surrogate Models ◽  
Robust Design Optimization ◽  
High Pressure Turbine ◽  
Surrogate Modelling ◽  
Multi Objective

To achieve reverse objectives in engine design, advanced modelling and analysis methods are among the key research technologies. In the presented work, a robust design optimization of a first stage high pressure turbine blade has been carried out. This blade derives from a current production of a Rolls-Royce aero engine. The motivation of this work is to show that the methodology of robust design optimization can be applied to high pressure turbine blades. A fully automated workflow, which encapsulated the integral blade design and analysis process, has been used. The main workflow objective is a representative life value of the external surface of the blade. In addition, the workflow enables the engineering uses to consider sub objectives like mass, efficiency and life at critical locations of the blade. These can also be taken into account in the multi-objective robust design optimization. This research also focuses on the use of surrogate models, with attention to the delivery of a physically correct result. For this purpose, the validation of the applied methods has a huge significance and a toolbox was created to generate and evaluate the quality of the surrogate models. In the present case sixteen geometry parameters were considered. In order to show that this methodology is not limited to geometry variation, parameters for material specification and for boundary conditions were varied in addition. The surrogate model was trained by the workflow generated DoE-data and could be used for different kinds of optimization. As a conclusion, it has been demonstrated that the methodology can be used for the engineering design process of turbine blades, while delivering physically correct results. The different techniques for surrogate modelling were examined and compared. With the help of these surrogate models, an optimization of life, mass and efficiency with 22.5 million evaluations was possible. Finally, an overview of the methodology for the case of a real world turbine blade could be given, and an improved blade in the sense of multi-objective robust design was found.

Sign in / Sign up

Export Citation Format

Share Document