High Pressure Turbine Blade Research Articles

Numerical investigation of high pressure turbine blade tip-shaping effects on the aerothermal and dynamic performance

Purpose In order to improve the engine reliability and efficiency, an effective way is to reform the turbine blade tip conformation. The paper aims to discuss this issue. Design/methodology/approach The present research provides several novel tip-shaping structures, which are considered to control the blade tip loss. Four different tip geometries have been studied: flat tip, squealer tip, flat tip with streamwise ribs and squealer tip with streamwise ribs. The tip heat transfer and leakage flow are both analyzed in detail, for example the tip heat transfer coefficient, tip flow and local pressure distributions. Findings The results show that the squealer seal and streamwise rib can reduce the tip heat transfer and leakage loss, especially for the squealer tip with streamwise ribs. The tip and near-tip flow patterns at the different locations of axial chord reflect that both the squealer seal and streamwise rib structure can control the tip leakage flow loss. In addition, the analysis of the aerodynamic parameters (the static pressure and turbine efficiency) also indicates that the squealer tip with streamwise ribs obtains the highest adiabatic efficiency with an increase of 2.34 percent, compared with that of the flat tip case. Originality/value The analysis of aerothermal and dynamic performance can provide a reference for the blade tip design and treatment.

Accuracy, efficiency and scalability of explicit and implicit FR/CPR schemes in large eddy simulation

The accuracy, efficiency and scalability of explicit and implicit temporal schemes associated with FR/CPR spatial schemes are studied in the context of large eddy simulation. One low pressure turbine and one high pressure turbine blade cases with different Mach numbers and Reynolds numbers are employed in this study. The 3 stage SSP Runge–Kutta (RK) scheme is compared with implicit backward difference formulas (BDF) of first and second order accuracy with a non-linear LU-SGS solver in the present evaluation. Various factors such as inner convergence tolerance, the frequency of the implicit operator update, and the order of time accuracy are investigated for large eddy simulation. The implicit BDF2-LUSGS algorithm can achieve good time accuracy, and is more efficient than the 3rd order explicit RK scheme, but not as scalable on a small cluster. At a very low Mach number, the explicit scheme is clearly not adequate, suggesting the need for pre-conditioning and/or an implicit scheme.

Effect of in-service burnout on the transonic tip leakage flows over flat tip model

Un-shrouded turbine blades are more common than shrouded ones in gas turbine aero-engines since they reduce the weight and avoid the centrifugal loading caused by the blades’ shrouds. Despite these important advantages, the absence of the shroud leads to leakage flows across the tip gap and exposes the blade tip to high thermal load and thermal damages. In addition, the leakage flows can contribute up to 30% of the aerodynamic loss in a turbine stage. In this study, the effect of in-service burnout is explored using a fundamental flat tip model of a high-pressure gas turbine blade. This investigation is carried out both experimentally in a transonic wind tunnel and computationally using the Reynolds Averaged Navier-stokes approach at high-speed conditions. It is found that exposing the tip to the in-service burnout effect changes the leakage flow behaviour significantly when compared with the tip with sharp edges (i.e. the tip at the start of its operational life). Different flow acceleration, flow structure and shockwave pattern and interactions are captured for the round-edge flat tip (i.e. the tip exposed to in-service burnout). The effective tip gap is found to be much larger for the round-edge flat tip allowing more leakage flow into the tip gap which results into higher tip leakage losses in comparison to the sharp-edge tip. Experimental and computational flow visualisations, surface pressure distributions and discharge coefficient are given and analysed for several pressure ratios over the tip gap.

МЕТОДЫ КОНТРОЛЯ ВНУТРЕННЕЙ ПОЛОСТИ ОХЛАЖДАЕМЫХ ЛОПАТОК ТВД С ПЕРСПЕКТИВНЫМИ СИСТЕМАМИ ОХЛАЖДЕНИЯ

The cooled turbine blades are very complex, labor-intensive parts, both for manufacturing and for inspection. With the development of aviation gas-turbine engines, the design of cooled turbine blades is becoming more complicated. To create new promising cooled turbine blades requires the introduction of new advanced manufacturing technologies and new progressive methods of dimensional control. This paper discusses the existing methods of inspection, which are used in the manufacture of turbine blades in mass production. Their advantages and disadvantages are given. The results of the work carried out in the framework of the international project "AERO.UA" are presented. The work was carried out to develop the inspection method of the internal cavity of the high-pressure turbine (HPT) blade with a promising cooling system developed at Ivchenko-Progress. The inspection of the internal cavity of the HPT blade with a promising cooling system was performed using computed tomography. Computed tomography performed at the Air Force Institute of Technology (ITWL) Warsaw on the General Electric X-Ray phoenix V | tome | X m tomograph. According to the results of testing the inspection method, the possibility of identifying manufacturing defects, the detection of which under the conditions of steady-state production is associated with very labor-intensive operations, is shown, and in some cases, inspection is impossible at all without destroying the part or performing expensive repairs. An assessment was made of the possibility of monitoring the geometrical dimensions of the internal cavity and external surfaces of cooled turbine blades. With the use of computed tomography, the dimensional control is possible with an accuracy of 0.01 mm. But to bind the blades to the base surfaces specified in the drawing when scanning, you must use a special snap. To confirm the possibility of controlling the dimensions specified in the drawing, it is necessary to continue work on scanning the cooled turbine blades using special tools.

Multi-axis low-cycle creep/fatigue life prediction of high-pressure turbine blades based on a new critical plane damage parameter

To study the multi-axis low-cycle creep/fatigue life of high-pressure turbine blades, a multi-axis low-cycle fatigue life prediction model based on a new critical plane damage parameter is proposed in this paper. Firstly, based on the thermal-structural coupling analysis of high-pressure turbine blades, the law of stress-strain distribution under the complex load of centrifugal load, temperature load and the aerodynamic load is obtained. Secondly, the multi-axis low-cycle fatigue life prediction model and the L-M equation are applied to predict the fatigue life and creep life of blade respectively. Furthermore, different experimental loading schemes are used to study the fatigue damage mechanism for the creep/fatigue interaction under high temperature condition, and a creep/fatigue interaction life prediction model is simultaneously provided. And finally, the experiment results demonstrate the rationality of the multi-axis low-cycle fatigue life prediction model and the creep/fatigue interaction should not be ignored.

Design and analysis of a high-pressure turbine blade in a jet engine using advanced materials

In this study, the use of Inconel 625, Palladium alloy and Titanium alloy have been done in jet turbine blade. The modelling of the turbine blade was done in Solidworks 16.0 software and its simulation was done under numerical computational method using Ansys 16.0 software. The tests conducted in the software were the equivalent stress (von misses stress), equivalent strain, and total deformation. Three types of material configuration were tested and compared: using only Ni-based alloy, using Ni-based alloy with Titanium and using Ni-based alloy along with Pd-alloy at high stress zones of the blade.

Rib Turbulator Heat Transfer Enhancements at Very High Reynolds Numbers

High-pressure stage gas turbine blades feature serpentine passages where rib turbulators are installed to enhance heat transfer between the relatively colder air bled off from the compressor and the hot internal walls. Most of the prior studies have been restricted to Reynolds number of 90,000 and several studies have been carried out to determine geometrically optimized parameters for achieving high levels of heat transfer in this range of Reynolds number. However, for land-based power generation gas turbines, the Reynolds numbers are significantly high and vary between 105 and 106. The present study is targeted toward these high Reynolds numbers where traditional rib turbulator shapes and prescribed optimum geometrical parameters have been investigated experimentally. A steady-state liquid crystal thermography technique is employed for measurement of detailed heat transfer coefficient. Five different rib configurations, viz., 45 deg, V-shaped, inverse V-shaped, W-shaped, and M-shaped have been investigated for Reynolds numbers ranging from 150,000 to 400,000. The ribs were installed on two opposite walls of a straight duct with an aspect ratio of unity. For very high Reynolds numbers, the heat transfer enhancement levels for different rib shapes varied between 1.4 and 1.7 and the thermal hydraulic performance was found to be less than unity.

An Error Indicator-Based Adaptive Reduced Order Model for Nonlinear Structural Mechanics—Application to High-Pressure Turbine Blades

The industrial application motivating this work is the fatigue computation of aircraft engines’ high-pressure turbine blades. The material model involves nonlinear elastoviscoplastic behavior laws, for which the parameters depend on the temperature. For this application, the temperature loading is not accurately known and can reach values relatively close to the creep temperature: important nonlinear effects occur and the solution strongly depends on the used thermal loading. We consider a nonlinear reduced order model able to compute, in the exploitation phase, the behavior of the blade for a new temperature field loading. The sensitivity of the solution to the temperature makes the classical unenriched proper orthogonal decomposition method fail. In this work, we propose a new error indicator, quantifying the error made by the reduced order model in computational complexity independent of the size of the high-fidelity reference model. In our framework, when the error indicator becomes larger than a given tolerance, the reduced order model is updated using one time step solution of the high-fidelity reference model. The approach is illustrated on a series of academic test cases and applied on a setting of industrial complexity involving five million degrees of freedom, where the whole procedure is computed in parallel with distributed memory.

Evaluation of microstructural degradation in a failed gas turbine blade due to overheating

Gas turbine blades may undergo overheating, which could cause serious microstructural degradation and even failure of turbine components. Nonetheless, limited published investigations focus on the evaluation of the microstructural degradation in overheated turbine blades. In this study, a high-pressure turbine blade was investigated, which comprised of equiaxed-cast superalloy substrate and an AlSi coating. The blade failed due to material loss at the airfoil tip of the leading edge. A systematic microstructural investigation of the failed blade was conducted, and the results were compared with those of thermally exposed samples. By taking several microstructural degradation parameters as references, the microstructural degradation of the failed blade was evaluated and the equivalent maximum service temperature was estimated. The results indicated that most locations of the failed blade had been operating under normal conditions. However, the airfoil tip of the leading edge had suffered serious overheating, and the overheating temperature was almost 200 °C above the normal service temperature. The overheating led to incipient melting of the substrate and material loss of the serviced blade. This study provides a guidance on the evaluation of microstructural degradation and failure analysis in conventionally cast turbine blades.

A framework for fatigue reliability analysis of high-pressure turbine blades

Fatigue evolution under continued stresses is a process of degradation of material performance with many uncertainties. In order to quantify the uncertainties of materials and working conditions, a probabilistic method is utilized to estimate the reliability of structures by considering scatter of the fatigue life prediction model, in which improvements are provided to model the accumulation of the damage. Firstly, the fatigue parameters are modeled by the Bayesian theory and the finite element analysis. Secondly, the distributions of parameters are transformed by the probabilistic method into the distribution of fatigue life by using the fatigue life prediction model, and a damage accumulation model is chosen to characterize regulation evolution of properties. Finally, the probability distribution function transformation approach is employed to expound distribution of fatigue damage by the known distribution of fatigue life, and a general probabilistic method is then used to estimate the reliability. By combining the above methods, the framework for reliability analysis is established and then is used to calculate the reliability for high-pressure turbine blades in a low cycle fatigue region under variable amplitude loadings.

Computational Fluid Dynamics Investigation of the Flow of Trailing Edge Cooling Slots

Slot film cooling is a popular choice for trailing edge (TE) cooling in high pressure (HP) turbine blades because it can provide more uniform film coverage compared to discrete film cooling holes. The slot geometry consists of a cutback in the blade pressure side connected through rectangular openings to the internal coolant feed passage. The numerical simulation of this kind of film cooling flows is challenging due to the presence of flow interactions such as step flow separation, coolant-mainstream mixing, and heat transfer. The geometry under consideration is a cutback surface at the trailing edge of a constant cross-section aerofoil. The cutback surface is divided into three sections separated by narrow lands. The experiments are conducted in a high-speed cascade in Oxford Osney Thermo-Fluids Laboratory at Reynolds and Mach number distributions representative of engine conditions. The capability of computational fluid dynamics (CFD) methods to capture these flow phenomena is investigated in this paper. The isentropic Mach number and film effectiveness are compared between CFD and pressure sensitive paint (PSP) data. When compared with the steady k − ω shear stress transport (SST) method, scale adaptive simulation (SAS) can agree better with the measurement. Furthermore, the profiles of kinetic energy, production, and shear stress obtained by the steady and SAS methods are compared to identify the main source of inaccuracy in RANS simulations. The SAS method is better to capture the unsteady coolant–hot gas mixing and vortex shedding at the slot lip. The cross flow is found to affect the film significantly as it triggers flow separation near the lands and reduces the effectiveness. The film is nonsymmetric with respect to the half-span plane, and different flow features are present in each slot. The effect of mass flow ratio (MFR) on flow pattern and coolant distribution is also studied. The profiles of velocity, kinetic energy, and production of turbulent energy are compared among the slots in detail. The MFR not only affects the magnitude but also changes the sign of production.

Entropy Analysis of the Flat Tip Leakage Flow with Delayed Detached Eddy Simulation.

In unshrouded turbine rotors, the tip leakage vortices develop and interact with the passage vortices. Such complex leakage flow causes the major loss in the turbine stage. Due to the complex turbulence characteristics of the tip leakage flow, the widely used Reynolds Averaged Navier–Stokes (RANS) approach may fail to accurately predict the multi-scale turbulent flow and the related loss. In order to effectively improve the turbine efficiency, more insights into the loss mechanism are required. In this work, a Delayed Detached Eddy Simulation (DDES) study is conducted to simulate the flow inside a high pressure turbine blade, with emphasis on the tip region. DDES results are in good agreement with the experiment, and the comparison with RANS results verifies the advantages of DDES in resolving detailed flow structures of leakage flow, and also in capturing the complex turbulence characteristics. The snapshot Proper Orthogonal Decomposition (POD) method is used to extract the dominant flow features. The flow structures and the distribution of turbulent kinetic energy reveal the development of leakage flow and its interaction with the secondary flow. Meanwhile, it is found that the separation bubble (SB) is formed in tip clearance. The strong interactions between tip leakage vortex (TLV) and the up passage vortex (UPV) are the main source of unsteady effects which significantly enhance the turbulence intensity. Based on the DDES results, loss analysis of tip leakage flow is conducted based on entropy generation rates. It is found that the viscous dissipation loss is much stronger than heat transfer loss. The largest local loss occurs in the tip clearance, and the interaction between the leakage vortex and up passage vortex promotes the loss generation. The tip leakage flow vortex weakens the strength of up passage vortex, and loss of up passage flow is reduced. Comparing steady and unsteady effects to flow field, we found that unsteady effects of tip leakage flow have a large influence on flow loss distribution which cannot be ignored. To sum up, the current DDES study about the tip leakage flow provides helpful information about the loss generation mechanism and may guide the design of low-loss blade tip.

Experimental and Numerical Investigation of the Flow in a Trailing Edge Ribbed Internal Cooling Passage

The flow field in a ribbed triangular channel representing the trailing edge internal cooling passage of a gas turbine high-pressure turbine blade is investigated via magnetic resonance velocimetry (MRV) and large eddy simulation (LES). The results are compared to a baseline channel with no ribs. LES predictions of the mean velocity fields are validated by the MRV results. In the case of the baseline triangular channel with no ribs, the mean flow and turbulence level at the sharp corner are small, which would correspond to poor heat transfer in an actual trailing edge. For the staggered ribbed channel, turbulent mixing is enhanced, and flow velocity and turbulence intensity at the sharp edge increase. This is due to secondary flow induced by the ribs moving toward the sharp edge in the center of the channel. This effect is expected to enhance internal convective heat transfer for the turbine blade trailing edge.

A Comparison of Experimental and Computational Heat Transfer Results for a Leading Edge Impingement System

Leading edge impingement systems are increasingly being used for high pressure turbine blades in gas turbine engines, in regions where very high heat loads are encountered. The flow structure in such systems can be very complex and high resolution experimental data is required for engine-realistic systems to enable code validation and optimal design. This paper presents spatially resolved heat transfer distributions for an engine-realistic impingement system for multiple different hole geometries, with jet Reynolds numbers in the range of 13,000–22,000. Following this, Reynolds-averaged Navier-Stokes computational fluid dynamics simulations are compared to the experimental data. The experimental results show variation in heat transfer distributions for different geometries, however average levels are primarily dependent on jet Reynolds number. The computational simulations match the shape of the distributions well however with a consistent over-prediction of around 10% in heat transfer levels.

High temperature materials for aerospace applications: Ni-based superalloys and γ-TiAl alloys

Abstract Two outstanding aerospace-oriented high-temperature materials, the single-crystal nickel-based superalloys for high-pressure turbine blades and the γ-TiAl-based alloys for low pressure turbine blades, are being presented here. In both cases, the optimisation of their mechanical properties is based on a high knowledge of metallurgy, mixing together different aspects such as processes, alloy design, deformation mechanisms, impact of oxidative environment or interaction between protective layers and protected alloy. Historical evolutions are recalled and put into perspective with more recent research activities.

Effects of Compressor Fouling and Compressor Turbine Degradation on Engine Creep Life Consumption

The effect of engine degradation in the form of compressor fouling and compressor turbine degradation on the creep life consumption of the high-pressure (HP) turbine blades of an LM2500+ industrial gas turbine engine is investigated in this work. The degradations are flow capacity degradation and isentropic efficiency degradation. An engine model was created in Cranfield gas turbine performance and diagnostics software, pythia. Blade thermal and stress models were developed together with the Larson–Miller parameter (LMP) method for creep life analysis. The percentage decreases in creep life due to each effect were examined. For the engine considered, compressor degradation has more impact on engine creep life toward peak power operation, while HP turbine degradation has more impact on creep life at lower power levels. The results of this work will give engine operators an idea of how engine components creep life is consumed and make reasonable decisions concerning operating at part loads.

Rapid Airfoil Design for Uncooled High Pressure Turbine Blades

Rapid Airfoil Design for Uncooled High Pressure Turbine Blades

CONVERTING THE AL-31ST ENGINE TO THE CONDITIONS OF MAIN GAS PIPELINES COMPRESSOR STATIONS

Turbojet engines designed for military and civil aviation are intended for flights at level from 9000 to 12000 m, where the air temperature is stable (minus 55 °C), and the air pres- sure corresponds to 150 mm Hg which is five times less than the atmospheric pressure at the surface of the Earth. It is shown that the AL-31ST gas turbine drive has a high sensitivity to the change in the temperature of the air entering the compressor. In the conditions of main gas pipelines compressor stations, in the territories with a continental and sharply continental climate, the air temperature varies greatly (for example, in the conditions of Bashkortostan during 24 hours, air temperature fluctuations can reach 18 °C), as a result of which the operation of the gas turbine engine is unstable, and the blades of high pressure turbine (HPT) experience overload and often break down. In the article, in order to determine the level of rational power loading of GPU-16R «Ufa», in the range of the boundary air temperatures with a deviation of ± 30 °C from the nominal value of 15 °C, at which the blade feather experiences the least load with bending moments, the AL-31ST drive has been recalculated to conditions of Earth-surface application. The analysis of differences between operating and nominal parameters is performed. A flow model of a rotor blade in the relative motion of a gas flow is created and the results of calculation in the form of velocity and pressure fields in the COMSOL Multiphysics 4 program are presented for the air temperature range from minus 15 oС to 45 oС. A signifi-cant difference in the modes is revealed, and the fact of HPT blades overload with bending moments is established, up to 5 % at the boundary temperatures. As a result, the hypothesis of specialists investigating the causes of breakage of the HPT blades was confirmed, that the breakage of the blades is associated with «... related fac-tors», and not «with defects of a metallurgical or casting nature». The HPT blade breaks from the material fatigue caused by periodic bending moments due to fluctuations in the outside air temperature, exceeding the permissible power load up to 5 % in the operating temperature range. Based on the analysis of the AL-31ST drive operation in the range of the outside air tem-perature limits (± 30 oС from the nominal temperature), it is recommended to maintain the operating mode of this drive of no more than 0.95 of the nominal load power.

Analysis of transient radiation-convection coupled effects on the HP turbine blade heat load with hot streak inlet

Hot Streak (HS) and thermal radiation are the two important factors affecting the heat transfer process in High Pressure (HP) turbine stage, while the previous investigations were limited in studying either of them independently. In order to investigate the combined effects of the HS and radiation, the present study is aimed at unsteady convection-radiation coupled simulations with regards to HS and uniform inlet conditions. Three dominant radiative factors were analyzed in details. Compared with the uniform cases, it is found that there are two high temperature regions in the rotor passages of HS cases. The maximum temperature increase of the vane surface is about 40 K caused by radiation. As the HS changes the temperature distribution, the radiative heat flux is redistributed and a portion of radiative heat flux transfer from the stator passage to the rotor passage. The turbine inlet radiation mainly affects on the guide vane, whereas the gas radiation has a more significant influence on the rotor blade heat load.

AEROTHERMAL OPTIMIZATION OF SQUEALER GEOMETRY IN AXIAL FLOW TURBINES USING GENETIC ALGORITHM

In turbomachines, a tip gap is required in order to allow the relative motion of the blade and to prevent the blade tip surface from rubbing. This gap which lay out between the blade tip surface and the casing, results in fluid leakage due to the pressure difference between the pressure side and the suction side of the blade. The tip leakage flow causes almost one third of the aerodynamic loss and unsteady thermal loads over the blade tip. Previous experimental and numerical studies revealed that the squealer blade tip arrangements are one of the effective solutions in increasing the aerothermal performance of the axial flow turbines. In this paper the tip leakage flow is examined and optimized with the squealer geometry as a means to control those losses related with the tip clearance. The squealer height and width have been selected as design parameters and the corresponding computational domain was obtained parametrically. Numerical experiments with such parametrically generated multizone structured grid topologies paved the way for the aerothermal optimization of the high pressure turbine blade tip region. Flow within the linear cascade model has been numerically simulated by solving Reynolds Averaged Navier-Stokes (RANS) equations in order to produce a database. For the numerical validation a well-known test case, Durham cascade is investigated in end wall profiling studies has been used. Sixteen different squealer tip geometries have been modeled parametrically and their performance have been compared in terms of both aerodynamic loss and convective heat transfer coefficient at blade tip. Also, these two values have been introduced as objective functions in the optimization studies. A state of the art multi-objective optimization algorithm, NSGA-II, coupled with an Artificial Neural Network is used to obtain the optimized squealer blade tip geometries for reduced aerodynamic loss and minimum heat transfer coefficient. Optimization results are verified using CFD.