High Pressure Turbine Blade Research Articles

Introduction of Axisymmetric Grooves as a Tip Seal Treatment for Small-Core Turbines

Abstract The aerodynamic penalties associated with the tip gap flow in axial turbines remains a challenging problem for turbine manufacturers. As modern gas turbines with small-core architectures are brought online, the influence of the tip gap continues to grow. While technologies to reduce tip losses have been implemented into the blades, little attention has been paid to the casing around the rotor. In this study, an introduction of axisymmetric groove enhancements for the casing are examined computationally. These studies use the National Experimental Turbine (NExT) geometry, an engine-representative high-pressure turbine blade. Steady, RANS simulations are used to assess the basic characteristics of grooves, such as depth, location, and arrangement. The objective of this introductory study was to determine the feasibility of impacting the tip leakage vortex formation and the associated losses in rotor efficiency. Furthermore, analyses were done with different tip gap heights along with both flat- and squealer-tipped blades. Tip seals with a single groove are demonstrated to improve rotor aerodynamic efficiency relative to ungrooved seals by 0.4 points when applied to flat-tipped rotor blades and 0.15 points with squealer tips. Arrays of grooves show improvements for flat-tipped blade performance by 0.76 points, while having little additional aerodynamic effect on squealer tip over single-groove designs. Finally, grooved tip seals appear to exert greater influence on turbine aerodynamic performance at larger tip gaps, indicating that grooved tip seals alter the sensitivity of rotor performance to the tip clearance gap.

CFD simulation and aerodynamic optimization of two-stage axial high-pressure turbine blades

CFD simulation and aerodynamic optimization of two-stage axial high-pressure turbine blades

Numerical simulation of the creep of a turbine blade made of a monocrystalline alloy

The article is devoted to the creep model of a monocrystalline alloy and the development of a methodology for identifying material parameters based on the results of physical experiments. A finite element analysis of the creep of a gas turbine engine blade was performed. Creep is one of the most dangerous types of deformation in the operating conditions of turbine blades. In the process of studying the problems of assessing the strength of turbine blades of aircraft engines and power plants, special attention should be paid to the study of stress redistribution during creep. The characteristics of the crystallographic structures of modern turbine blades have a very significant influence on the progress of the crack development process during engine operation. To date, turbine blades are manufactured by the method of single-crystal casting. This type of blade material structure is characterized by orthotropic mechanical properties. This study considers the steady-state creep model of an anisotropic heat-resistant single-crystal alloy with cubic symmetry. The authors carried out a numerical simulation of the material parameters using the well-known literary creep properties of single crystals. An algorithm is described that allows you to determine some creep characteristics of single crystals. The parameters of the given ratios can be obtained after conducting direct experiments, or based on micromechanical analysis, using the example of composite materials. The authors calculated the creep constants of a typical heat-resistant monocrystalline alloy as a result of the approximation of its creep curves, which were obtained experimentally. Based on the Norton-Bailey equation and using the Maple Release 2021.0 calculation complex, a graph of the dependence of the rate of creep deformation on the level of load applied to the material was plotted, and the minimum rate of deformation and creep constants were also determined. The results of the calculations were used for finite-element simulation of creep on the example of a solid-state model of a high-pressure turbine blade. Several series of calculations were carried out on the basis of the ANSYS Workbench complex, in particular, the calculation of the elastic problem when the blade is loaded by centrifugal forces, as well as the accumulation of creep deformations at different exposure times. Graphs of the change in equivalent stresses and creep deformations as a function of time are plotted.

Investigation of temperature assisted corrosion failure of an aircraft turbine blade

In this study, the root cause of a recurring corrosion problem at the tip of an internally cooled high-pressure turbine blade (HPT) of a turbofan engine was investigated. A two-pronged strategy was utilized involving (i) experimental techniques to identify the blade material and corrosion products, and (ii) computational methods to analyze the fluid flow over the blade. In the experimental thrust, an array of techniques was employed to determine the chemical composition, microstructure, corrosion products and microhardness values while the airflow, pressure distribution and thermal profile over the blade were found through computational paradigms. The blade was a Ni-based superalloy containing typical gamma (γ) and gamma prime (γ′) phases, with an average microhardness value of 389 HV at room temperature. SEM analysis of the corroded area revealed oxides of iron, calcium, magnesium, aluminium, and silicon (FCMAS) most of which were not part of the base alloy. To determine the temperature profile and pressure distribution on the blade, a conjugate heat transfer analysis using Ansys CFX was conducted. A combination of experimental data from the engine ground running test station and analytical equations was used to find the required boundary conditions for the computational analysis. Pressure distribution over the entire blade was found which showed higher flow velocities and cross flows at the blade tip which potentially cause erosion at that region in the presence of impurities in the incoming air. Temperature distribution over the entire blade was also obtained, and it was found that the highest temperatures for the internally cooled blade occurred at the tip region which were in the range of ∼1060 °C–1250 °C (average ∼1114 °C). The temperature on the tip was high enough to cause the melting of the impurities that ingress the coating and cause its degradation during thermal cycling. This exposes the base alloy to FCMAS and causes corrosion. These high temperatures (1060–1250 °C) also induced phase transformation and increased the brittleness (∼22 % increase in hardness) which aggravated the presence of corrosion and resulted in crack formation and blade rejection.

Research on the flow and heat transfer performance of variable geometry high-pressure turbine blade tip area of gas turbine

Research on the flow and heat transfer performance of variable geometry high-pressure turbine blade tip area of gas turbine

Resource-efficient regeneration by generating a digital component image based on different NDT techniques

Manufacturing of components that are subjected to high stress levels during operation, such as aircraft engine turbine blades and battery calender rollers, is resource intensive and costly. These components are often designed not as a mono-material, but as a layered system. If operational stresses initiate crack-like defects or changes in the microstructure in the coating and/or base material, this can lead to premature component failure. To prevent this, the component’s condition is assessed during maintenance by inspections. Based on the data obtained, a decision is made whether the component can remain in operation, must be replaced or needs to be repaired. This decision is based on safety, economic and environmental considerations. In order to repair or regenerate such capital goods in a targeted and resource-efficient manner, a differentiated assessment of surface defects and material changes in the coating as well as in the base material is required. This can be achieved by an intelligent combination of different non-destructive testing techniques. In this paper, the differentiated assessment of the condition of the individual layers and the base material is presented using the example of a high-pressure turbine blade. The test results of the non-destructive testing techniques used are combined in a digital image of the component from which suitable repair measures can be derived. The transfer of the findings from the turbine blade to the layered system of a calender roller is also discussed.

Simultaneous improving high temperature oxidation resistance and room temperature toughness of Nb[sbnd]Si based alloy via appropriate Ta content addition strategy

NbSi based alloys are expected to become the application materials for high pressure turbine blades of aircraft engines. This paper investigates the impact of Ta addition on room temperature fracture toughness and isothermal oxidation behavior at 1250 °C of four NbSi based alloys with compositions of Nb-16Si-20Ti-8Zr-4Hf-1.5Al-xTa (x = 0, 1, 2, and 4 at.%) fabricated through vacuum non-consumable arc-melting. The results show that the room temperature fracture toughness and 1250 °C oxidation resistance of the alloys increase first and then decrease with the addition of Ta element. The room temperature fracture toughness of 1Ta alloy is the best, reaching 17.1 MPa·m1/2. Adding a moderate amount of Ta (1 at.%) significantly enhances the alloy's oxidation resistance, but an excessive Ta content diminishes this enhancement. This research has obtained a Nb-16Si-20Ti-8Zr-4Hf-1.5Al-1Ta alloy with high room temperature fracture toughness and excellent high temperature oxidation resistance simultaneously, which provide an ideal path for designing NbSi alloys with superior comprehensive properties.

Effects of blowing ratios and lip thicknesses on film cooling for trailing edge cutback with latticework ducts

In this study, a trailing edge cutback configuration with latticework ducts is proposed to provide more effective thermal protection for high-pressure turbine blades. Detached eddy simulation, as a LES/RANS hybrid method, is utilized to investigate the film cooling efficiency and turbulent behavior of the trailing edge cutback. The flow dynamics mechanism of cutback film cooling is elucidated by analyzing five blowing ratios (0.2 ∼ 1.25) and four lip thicknesses (0.5 ∼ 2) schemes. Time-averaged results indicate that the counter-intuitive efficiency decline and recovery behavior is identified again. Quantitatively, the area-averaged film cooling efficiency decreases by 14.3 % in the efficiency decline interval and increases by 16.3 % during the recovery interval. The distinct deflection of the swirling coolant jet leads to an efficiency distribution characterized by higher values on the sides and lower values in the center. Moreover, the mixing of the swirling jets and crossflow produces the Asymmetric Counter-rotating vortex pairs and Spiral-vortex system, which exacerbate heat transfer near the wall. At the highest blowing ratio, the excess coolant is forced against the wall by a rotating streamwise vortex system, providing better film coverage. Additionally, the film cooling efficiency of the cutback structure with a lip-to-slot ratio of 2 decays by 27 % as the blowing ratio decreases. The detrimental influence of the lip thickness increase on the film cooling efficiency is considerably mitigated by the swirling jet, especially in the maximum lip thickness.

Restoration of high-pressure turbine blade combs using a concentrated energy flow

A useful model of temperature distribution in the deposited layer of ВЖ159 powder and in the zone of its connection with the base material — ЖС6У alloy, during laser cladding is developed. The optimal modes of laser cladding of powder onto alloy blades are determined. The properties of the deposited layer and the quality of its bonding with the base material are studied. It is shown that there are no defects such as pores, cracks and discontinuities in the deposited layer and the interface zone. Keywords:restoration, laser cladding, blades, high-pressure turbine, temperature distribution, microstructure, microhardness, X-ray spectral microanalysis

A study of angle parametric optimization of the composite honeycomb on turbine blade tip

To reduce the secondary flow loss of the high-pressure turbine caused by the tip leakage flow, the design of the composite honeycomb tip has been experimentally and numerically examined in this study. The arrangement angle of the composite honeycomb is defined and optimized using the radial basis function and genetic algorithm. The tip flow field of the high-pressure turbine blade with a 2.128 mm clearance is simulated by CFX 18.0, and the arrangement angle of the composite honeycomb is optimized to obtain a lower loss. Moreover, the flat tip, basic honeycomb tip, and optimized honeycomb tip are tested in a low-speed cascade test facility. The results show that the vortices and radial velocity induced by the honeycomb structure cavity increase the kinetic energy loss of the leakage flow and reduce the leakage flow rate. Compared with the basic honeycomb, the optimized honeycomb has a supplemental blocking effect on the tip leakage flow by reducing the crosswise pressure gradient in the blade tip. The optimized honeycomb also changes the outlet angle of the leakage flow in the tip clearance, reduces the leakage vortex, and the total pressure loss is further reduced by 2.5%.

Unsteady Flows and Component Interaction in Turbomachinery

Unsteady component interaction represents a crucial topic in turbomachinery design and analysis. Combustor/turbine interaction is one of the most widely studied topics both using experimental and numerical methods due to the risk of failure of high-pressure turbine blades by unexpected deviation of hot flow trajectory and local heat transfer characteristics. Compressor/combustor interaction is also of interest since it has been demonstrated that, under certain conditions, a non-uniform flow field feeds the primary zone of the combustor where the high-pressure compressor blade passing frequency can be clearly individuated. At the integral scale, the relative motion between vanes and blades in compressor and turbine stages governs the aerothermal performance of the gas turbine, especially in the presence of shocks. At the inertial scale, high turbulence levels generated in the combustion chamber govern wall heat transfer in the high-pressure turbine stage, and wakes generated by low-pressure turbine vanes interact with separation bubbles at low-Reynolds conditions by suppressing them. The necessity to correctly analyze these phenomena obliges the scientific community, the industry, and public funding bodies to cooperate and continuously build new test rigs equipped with highly accurate instrumentation to account for real machine effects. In computational fluid dynamics, researchers developed fast and reliable methods to analyze unsteady blade-row interaction in the case of uneven blade count conditions as well as component interaction by using different closures for turbulence in each domain using high-performance computing. This research effort results in countless publications that contribute to unveiling the actual behavior of turbomachinery flow. However, the great number of publications also results in fragmented information that risks being useless in a practical situation. Therefore, it is useful to collect the most relevant outcomes and derive general conclusions that may help the design of next-gen turbomachines. In fact, the necessity to meet the emission limits defined by the Paris agreement in 2015 obliges the turbomachinery community to consider revolutionary cycles in which component interaction plays a crucial role. In the present paper, the authors try to summarize almost 40 years of experimental and numerical research in the component interaction field, aiming at both providing a comprehensive overview and defining the most relevant conclusions obtained in this demanding research field.

Adaptive directed support vector machine method for the reliability evaluation of aeroengine structure

Machine learning methods have been widely applied to structural reliability analysis, due to the excellent performance in modeling precision and efficiency. In this paper, an adaptive directed support vector machine model (ADSVM) is proposed by integrating the adaptive sampling technology (AST) and the developed directed support vector machine (DSVM), to improve the efficiency and accuracy of turbine blade reliability analysis. In the developed method, the AST is adopted to extract high quality samples, in respect of distributed probability density and the distances between sample values and response values. The DSVM was developed by introducing the penalty coefficients for all the slack variables, to improve the modeling accuracy of SVM approach. One numerical example and the reliability analysis of aeroengine high-pressure turbine blade were performed to validate the developed methods. As illustrated in this numerical investigation, the modeling precision of DSVM is improved by alleviating the lag effect of support vector regression. From the comparison of methods, it is demonstrated that the ADSVM has higher accuracy, efficiency and stability in reliability simulations, which are improved by ∼20 % and ∼46.2 %, respectively. The main efforts of this study are to propose a promising approach, ADSVM, for the reliability analysis of complex structures besides turbine blades, which hold the academical significance in enriching and developing mechanical reliability theory.

Multidisciplinary Automation in Design of Turbine Vane Cooling Channels

In the quest to enhance the efficiency of gas turbines, there is a growing demand for innovative solutions to optimize high-pressure turbine blade cooling. However, the traditional methods for achieving this optimization are known for their complexity and time-consuming nature. We present an automation framework to streamline the design, meshing, and structural analysis of cooling channels, achieving design automation at both the morphological and topological levels. This framework offers a comprehensive approach for evaluating turbine blade lifetime and enabling multidisciplinary design analyses, emphasizing flexibility in turbine cooling design through high-level CAD templates and knowledge-based engineering. The streamlined automation process, supported by a knowledge base, ensures continuity in both the mesh and structural simulation automations, contributing significantly to advancements in gas turbine technology.

Genetic Algorithm-Based Optimisation of a Double-Wall Effusion Cooling System for a High-Pressure Turbine Nozzle Guide Vane

Double-Wall Effusion Cooling schemes present an opportunity for aeroengine designers to achieve high overall cooling effectiveness and convective cooling efficiency in High-Pressure Turbine blades with reduced coolant usage compared to conventional cooling technologies. This is accomplished by combining impingement, pin-fin and effusion cooling. Optimising these cooling schemes is crucial to ensuring that cooling is achieved sufficiently at high-heat-flux regions and not overused at low-heat-flux ones. Due to the high number of design variables employed in these systems, optimisation through the use of Computational Fluid Dynamics (CFD) simulations can be a computationally costly and time-consuming process. This study makes use of a Low-Order Flow Network Model (LOM), developed, validated and presented previously, which quickly assesses the pressure, temperature, mass flow and heat flow distributions through a Double-Wall Effusion Cooling scheme. Results generated by the LOM are used to rapidly produce an ideal cooling system design through the use of an Evolutionary Genetic Algorithm (GA) optimisation process. The objective is to minimise the coolant mass flow whilst maintaining acceptable metal cooling effectiveness around the external surface of the blade and ensuring that the Backflow Margin for all film holes is above a selected threshold. For comparison, a Genetic Aggregation model-based optimisation using CFD simulations in ANSYS Workbench is also conducted. Results for both the reduction of coolant mass flow and the total optimisation runtime are analysed alongside those from the LOM, demonstrating the benefit of rapid low-order solving techniques.

Numerical study of a novel cooling protection scheme with rail crown holes for the squealer tip in a turbine blade

The squealer tip is acknowledged as an effective and dependable design for minimizing leakage loss and reducing thermal load in high-pressure turbine blades. After confirming the numerical approach, this study explored the cooling and aerodynamic characteristics of a novel cooling protection scheme with rail crown holes in a squealer tip. The rail crown hole parameters including the hole number, size, and distribution are research variable. Evaluation indexes of cooling and aerodynamic performance are the tip surface adiabatic film cooling efficiency (η) and clearance leakage flow rate (LFR). In cooling aspects, increasing the hole number or the hole size can improve the coolant attachment to the rail crown surface under the same coolant mass flow rate (Q). The coolant distribution within the cavity is substantially improved by concentrating the film holes at the leading-edge rail, which enhances the cooling protection of the cavity floor. In aerodynamic aspects, at low Q conditions, the total LFR correlates only with Q and is less sensitive to hole parameters. At high Q conditions, enlarging the hole size proves more effective in suppressing total LFR. Additionally, three cases with optimal cooling effects are chosen to investigate the impact of Q. These three cases are the scheme with an increasing hole number (case 1), the scheme with an enlarging hole size (case 4), and the scheme with concentrated holes at the leading edge (case 5). The results show that case 5 consistently exhibits superior cooling protection for the cavity floor in all Q conditions. For average η of the rail crown surface, cases 1 and 5 reach the peak value of average η at Q = 1.0Q0, while case 4 attains its peak value at Q = 1.5Q0.

The mechanical characteristics and cooling performance of a turbine blade with non-homogeneous lattice structure

The turbine blade is working at a high rotation speed and under constant impact from the high-temperature and high-pressure mainstream, causing deformation and a reduction in working efficiency; at the same time, the mainstream temperature at the inlet is increasing to achieve a high heat efficiency. Therefore, there is a high demand for turbine blade cooling technology. In this study, the authors proposed replacing the traditional spoiler with a non-homogeneous lattice structure in the inner cavity of the high-pressure turbine blade. The turbine blade was then simulated and experimend in real-world conditions using numerical simulation software and manufactured actual blades, and the maximum deformation of the turbine blade was measured to assess the impact of the non-homogeneous lattice structure on the blade's strength. The findings are as follows: The maximum deformation of the turbine blade optimized with a non-homogeneous lattice structure is 9.91% less than that of the traditional turbine blade with a spoiler, and its deformation-weight ratio is 11.51% less than the latter; the optimized turbine blade has a large cooling area in the concave surface of the blade, which improves the air film cooling effect, the coverage area of gas film hole outlet flow rate has increased by 12.39%. Finally, the non-homogeneous lattice structure optimizes the inner cavity structure of the turbine blade and is useful for improving the turbine blade's structural design and air film cooling performance.

Aerothermal performance of cavity tip with flow structure effects in a transonic high-pressure turbine blade

Understanding the mechanisms of leakage flow control and heat load distribution is important in optimizing the structure of cavity tips in gas turbines. This paper investigates the aerothermal performance of cavity tip in a transonic high-pressure turbine stage. By analyzing flow structure, the effects of squealer height, squealer width and gap height on aerothermal performance are explained. It has been found that both the leakage flow and heat load distribution are influenced by the structure of vortices in the cavity. The leakage flow is controlled by the suction-side cavity vortex and the scraping vortex. The heat load distribution relates to the “M-shaped” leakage flow induced by the scraping vortex. The geometric parameters can affect the structure of vortices, which then impacts the aerothermal performance of cavity tip. As the squealer height is 2.5τ0, both the leakage flow rate and average heat transfer coefficient decrease by 0.64 % and 6.6 %, respectively. A multi-objective optimization with a feedforward neural network prediction is conducted to determine the distribution characteristics of optimal design parameters. The optimal values are approximately 2.34 for the expansion ratio, 0.5τ0 for the squealer width and 0.5τ0 for the gap height. The optimal squealer height ranges from 1.5τ0 to 2.5τ0.

Optimization of a High Pressure Turbine Blade and Sector-Based Annular Rig Design for Supercritical CO2 Power Cycle Representative Testing

Abstract As part of the ongoing research into the design of hardware for zero emission cycles, a first-stage high-pressure turbine (HPT) blade is optimized for a 300 MWe supercritical CO2 (sCO2) power cycle using the surrogate-assisted genetic algorithm optimizer in Numeca FINE/Design three-dimensional with objectives of increasing efficiency and decreasing heat load to the blade. Supercritical CO2 property tables are constructed from NIST REFPROP data for the condensable gas simulation in FINE/Turbo. A detailed mesh sensitivity study is performed for a baseline design to identify the proper-grid refinement and efficiently allocate resources for the optimization. Seventy design variables are selected for the initial population generation. Self-organizing maps are then used to focus the design variables on the most important ones affecting the objective functions. The optimization results in approximately 3000 three-dimensional Reynolds Averaged Navier Stokes simulations of different blade shapes with increases in efficiency of up to 0.85% and decreases in heat load of 14%. Families of blade shapes are identified for experimental testing in an annular rig at the Purdue Experimental Turbine Aerothermal Laboratory. A design to adapt the annular cascade for testing optimized geometries is introduced, which features eccentric radius sectors allowing for scaled-up geometries of sCO2 optimized blade profiles to be tested at design cycle representative conditions at high Reynolds numbers in dry air. Analysis into the effects of Reynolds number, working fluid, and geometric relations are presented to prove the efficacy of the test method.

A cooled turbine blade design and optimization method considering the cooling structure influence

This study introduces a multidisciplinary design methodology tailored for enhancing the performance of cooled turbine blades by amalgamating thermal and aerodynamic calculation modules. The approach is unique in terms of its integration of a multi-objective optimization platform, aimed at refining aerodynamic performance and gauging the heat transfer capabilities during the preliminary aerodynamic design phase. To accomplish this objective, a one-dimensional pipe-network calculation tool was incorporated into the thermal module to quickly evaluate the heat transfer performance of the blades under different conditions. This tool also provides more realistic film hole inlet boundary conditions essential for three-dimensional aerodynamic calculations. Implementing this platform in optimizing a high-pressure turbine blade revealed a Pareto-optimal front, comprising −η1 and η2 (representing optimization objectives for aerodynamic and heat transfer performance, respectively), showcasing a constrained relationship. Upon scrutinizing three optimization cases against the prototype, optimization case 1 demonstrates the most significant enhancements in aerodynamic performance, showing a 0.2015% improvement in aerodynamic efficiency relative to the prototype. Conversely, optimization case 3 displays a comparatively modest augmentation in aerodynamic performance but excels notably in heat transfer performance, showcasing a 7.61% reduction in the maximum temperature of the blade surface compared to the prototype. Through adept optimization strategies and meticulous variable selection, we maintained a relatively stable mainstream mass flow across the optimization cases (less than 0.05% variation). These findings underscore the efficacy of our multidisciplinary design approach for cooled turbine blades, promising efficiency improvements in current design practices and potential reductions in project duration.

Silica-based ceramic cores for high-pressure turbine airfoil blades in aircraft engines

Ceramic cores for high-pressure turbine airfoil blades in aircraft engines are characterised by very complicated shapes and the presence of small diameter holes and thin long grooves. In this paper, eight starting powder mixtures having different contents of fused silica, zirconium silicate, alumina and borosilicate glass were used for preparation of core material. The composition, which satisfies every demand for core application in the investment casting of high-pressure turbine blades, contains 69.0wt.% fused silica, 13.0 wt.% zirconium silicate, 12.0wt.% alumina and 6.0wt.% borosilicate glass. This material was characterised by a mechanical strength of 33.1MPa, a coefficient of thermal expansion of 2.56 ? 10?6 1/K, surface roughness of 1.8 ?m, shrinkage of less than 0.8% and an average pore size diameter of 2.7 ?m. The thin-walled ceramic cores were formed by the high-pressure injection moulding method, which required the selection of a thermoplasticiser, feedstock formulation and determination of the optimal processing conditions protecting the injected cores from defects and deformation. The post-shaping process of the cores included both water and thermal debinding, sintering, precision machining and dimension measurements.