Refurbishing superalloy components for gas turbines

Abstract

Since 1968, FVV have developed specific maintenance techniques to ensure the safe operation of turbine blades. Among the new methods employed or shortly to be used is the non–destructive testing of blades by optical and scanning electron microscopy, and this is described in the paper. New possibilities for predicting remaining blade life under various conditions of stress and temperature are considered, and new types of coating obtained using plasma spraying and hot isostatic pressing are discussed.

MST/97     

Field Experience of Platinum Aluminide Coated Turbine Blades

This paper describes the results obtained from an evaluation of several platinum aluminide coated first stage turbine blades returned from the field. The IN-738LC blades had accumulated from 3,900 to 27,500 service hours in Centaur (1) 50 industrial gas turbine engines, operating in a high temperature oxidizing environment. The coating performance and condition were assessed using optical and electron microscopy. The condition of the coating was correlated to blade operating temperatures, which were estimated using the gamma prime coarsening technique. The degradation mechanism of the coating, remaining coating life, and blade repairability were also addressed.     

Constitutive modelling and creep life prediction of a directionally solidified turbine blade under service loadings

Abstract

The high pressure turbine (HPT) blade of a heavy duty gas turbine operates under the interaction of complex aerodynamic, centrifugal and thermal loadings. The reliability of continuous working at elevated temperatures is a major limitation in service application of these materials. Therefore, it is essential to build the constitutive equations for predicting and analysing the creep deformation and creep lifetime of the blade. In present work, the creep deformation and lifetime of a HPT blade made of a nickel-based directionally solidified (DS) superalloy was numerically predicted. The θ-projection method was used to characterize the creep deformation of DZ125 under different temperatures and stress levels. The uniaxial equations of the θ-projection method were expanded into multi-axial form which was implemented into ABAQUS/UMAT (User MATerial subroutine) by an Euler method. A modified θ-projection method was employed to promote the adaptability of the original model to both steady state and transient temperature fields. Transient stress, strain and displacement distribution of the critical position inside the blade were obtained for service loading. The Larson–Miller parameter was employed to predict the creep lifetime of the blade. Simulation and results of the θ-projection method may also provide suggestions for the safety and life evaluation of HPT blade and other turbine blades.     

The relative durability of a conventional and a platinum-modified aluminide coating

One of the major maintenance problems for gas turbine engine operators is the degradation of high pressure turbine blades and vanes. The Australian Department of Defence was experimenting unacceptably high rejection rates of these components in one of its high performance engines as a result of severe hot corrosion attack. Accordingly a programme was initiated to find a more durable coating than the conventional aluminide coating supplied for the first-stage blades by the engine manufacturer. The coating chosen for the trial was a platinum-modified aluminide coating. Metallographic examination of both coating systems after 750 hours of service in an operational aircraft clearly established the superiority of the platinum-modified coating. This coating has now been adopted by the Australian Defence Force as the specified coating for first-stage blades in this engine.     

Microstructure and properties of alloy HP50–Nb: comparison of as cast and service exposed materials

Abstract

Spun cast alloy HP50–Nb tubing retired from service in a steam reformer after approximately 8 years of service had stress rupture properties suggesting very short remaining life. Microstructural examination, however, indicated little or no creep damage (cavitation) but did reveal extensive microstructural modification. The microstructure, mechanical properties, and stress rupture properties of the ex service material were determined and compared with virgin as cast material. The microstructural changes that occurred during service included the formation of phases rich in alloying elements such as chromium and niobium that would otherwise be expected to contribute to high creep strength. The creep life of the service exposed material was correspondingly short. The creep ductility of the service exposed material, however, was significantly higher than that of the as cast material. The implications of these findings for remaining life assessments of reformer furnaces using these materials are discussed.

MST/3207     

Metallografische Restlebensdauerbestimmung an geschmiedeten Turbinenschaufeln (Legierung Nimonic 108)

Metallographic Determination of Residual Life Time for a Forged Turbine Blade (alloy Nimonic 108) A Metallographic procedure to investigate creep voids in turbine blades by means of a reference micrograph set makes it possible to determine the degree of damage and the residual life. The efficiency of hot isostatic pressing (regeneration) for healing creep voids is judged in the same way.     

Improvement of grain structure and mechanical properties of a land based gas turbine blade directionally solidified with liquid metal cooling process

Abstract

The manufacturing process of a directionally solidified (DS) IN738LC turbine blade, produced with the liquid metal cooling (LMC) process was improved based on process modelling. The improvement involved varying the system dimensions in the baffle area and optimising the mould thickness, design and the withdrawal parameters. The grain structure of the DS blades produced exhibits a well defined <001>. texture with a few stray grains near the blade top compared to the previous design. Some blades were given to a two stage heat treatment followed by tension tests at 25 and 650°C as well as creep tests at 152 MPa/982°C and 340 MPa/850°C. The yield and tensile strength of improved DS blades were comparable to conventionally cast (CC) IN738LC blades, while the creep properties and the tensile elongation of the DS blades were significantly improved using the optimised LMC process. The LMC system is under more modifications to produce defect free single crystal turbine blades.     

Review of rejuvenation process for nickel base superalloys

Abstract

Combustion turbine components exposed to elevated temperatures and high stresses are subject to periodic replacement as a result of mechanical property degeneration to levels below that required for continued safe operation. In view of the very significant cost associated with replacement components there has been much interest in the rejuvenation of parts. This report reviews published literature on the rejuvenation of nickel base superalloys. The restoration of microstructure and properties to levels equivalent to the original material have been achieved with the use of reheat treatment alone or recovery cycles incorporating both hot isostatic pressing and reheat treatment. The majority of the literature considers the restoration of creep properties, although some success in the rejuvenation of fatigue properties has also been reported.     

Non destructive investigation of the condition of Gas Turbine Blades

In order to determine the remaining life of service exposed turbine blades it is necessary to characterize the degeneration of the microstructure of the base metal during service. Since turbine blades of industrial gas turbines are kept in complete stages, non-destructive inspection (NDI) is very attractive. Hence the goal of the investigation reported here was to evaluate a NDI technique able to detect microstructural changes of the base metal. It was found that single-stage replication (in combination with investigation in a Scanning Electron Microscope) is a relatively simple technique that fulfils all requirements. This technique can be used in-situ on uncoated buckets. For coated turbine blades local removal of the coating is necessary to perform base metal replication.     

Zur Bewertung von Schutzschichten gegen Heißgaskorrosion an Gasturbinenschaufeln

On the Assessment of Protective Coatings to combat Hot Gas Corrosion on Gas Turbine Blades . Turbine blades of aero-engines are normally protected from hot gas corrosion by means of aluminium diffusion coatings. The inherent negative results of these coatings – the resistance to corrosion gains with the growing thickness of the coating, but the strength generally suffers – calls for an optimum balance between resistance to hot gas corrosion and resistance to mechanical stresses for a maximally long life of the blades. Conventional criteria – chiefly that of the progress of corrosion – not being adequate for a reliable assessment of the potential gain, it is recommended that creep and fatigue tests using simultaneously superimposed corrosion be employed for evaluating protective coatings. Tests conducted at faithfully simulated service conditions on samples and blades in cast nickel material IN 100 show how increasing thickness will cause the fatigue strength and the ductility of the coating to be reduced at a more or less rapid rate. The tests suggest the conclusion that conventional coating parameters still need improving if the optimum balance between resistance to hot gas corrosion and resistance to mechanical stresses shall be gained. The still small number of tests conducted cannot provide conclusive evidence. This paper is nevertheless published in the hope that it will invite in-depth studies of the problem of high temperature protective coatings under the additional aspect of true-to-life mechanical stresses.     

Accelerated LCF‐creep experimental methodology for durability life evaluation of turbine blade

This paper proposes an accelerated low cycle fatigue (LCF)‐creep experimental methodology in laboratory to investigate the durability life of turbine blades. A typical mission profile of the turbine blade was obtained by means of rain flow counting method, considering both the actual flight condition and ground test data. Finite element analysis (FEA) was conducted to obtain the stress and temperature fields of turbine blade. A test system was constructed to conduct LCF‐creep experiments of turbine blades, simulating the stress and temperature distributions of critical section properly. LCF‐creep experiments of full‐scale turbine blades were performed under a trapezoidal loading spectrum. Experiment results showed that the durability life of turbine blade based on numerical method was longer than that based on this experimental methodology, even an order of magnitude. Furthermore, this experimental methodology helped to extend the service life of this blade safely, and its validity was verified in actual service condition.     

De-alloying and fatigue of high temperature alloys used for gas turbine blades

The laws governing high temperature oxidation and de-alloying of high temperature nickel toasted alloys with aluminium content of up to 2% and their effect on the fatigue resistance at high temperatures are investigated. An experimental and calculation procedure to predict the elements distribution curves in the surface layer of blades, implemented by means of an algorithm and a program for solution of direct and inverse diffusion problems under the arbitrary law of temperature changing, with mobile boundary and selected limiting conditions has been developed.

With the availability of sufficient data for a given alloy that characterize the diffusion and the influence of de-alloying on the longevity, the procedure allows a prediction of the remaining service life of the gas turbine blades.     

Cyclic operation of aero gas turbines – materials and component life implications

Abstract

Historically, the issues connected with the lifing of power generation gas turbine components have been very different from those associated with aero engines. Specifically, component lives in the power generation application have been dictated by creep and high cycle fatigue, whereas low cycle fatigue has been the driver for aero engines. However, developments in the design and usage of gas turbines within the respective industries have resulted in this distinction becoming increasingly blurred. This paper highlights recent advances in the materials technology, stress analysis and lifing of aero engine components, which are potentially relevant to industrial gas turbines. In particular, the development of complex constitutive equations for modelling plasticity and anisotropic creep are discussed, with particular reference to the behaviour of single crystal turbine blades. Moreover, developments in the methodologies used to estimate safe service lives for the components are considered. Specifically, a new lifing procedure, capable of accurately predicting component lives from plain specimen data alone, is discussed.     

New nickel base investment casting alloys IN 6201 and IN 6203

Abstract

Alloy IN 6201 satisfies the requirement for an alloy as strong as IN 792 with the corrosion resistance of IN 939, for investment cast blades in advanced industrial and marine gas turbines. Alloy IN 6203 was developed to optimise the combination of creep rupture strength and hot corrosion resistance available in directionally solidified blades for marine gas turbines. Following an outline of the basis for selecting the compositions of these alloys, data are provided on foundry characteristics, mechanical properties, corrosion resistance, alloy stability, and the result of trial casting of blades in IN 6201.

MST/904     

Degradation of MCrAlY Coating and Substrate Superalloy During Long Term Thermal Exposure

Reliability and service life of hot gas path components of a gas turbine are limited by the degree of coating and substrate material degradation that occurs during service. MCrAlY type coatings and Ni-base alloys are widely used in the industry. In this study, the effect of long term thermal exposure with and without stress on a blade material, U520 and on a MCrAlY coating has been investigated. Microstructural degradation of coating and U 520 material are presented as a function of time, temperature and stress. These results are compared with the degradation observed in service-exposed blades.     

PROTECTIVE COATINGS FOR AERO ENGINE HOT SECTION COMPONENTS

Aluminide coatings have been widely used in the aircraft industries for the protection of gas turbine engine hot section components against oxidation and/or hot corrosion. This paper considers modes of coating degradation under conditions of cyclic oxidation, hot corrosion and corrosion-erosion interactions during service, as well as the effects of interdiffusion between coating and substrate alloys either during service or coating application. It also discusses means of improving existing coatings as well as advanced coating systems currently under development. In assessing coating performance, consideration must be given to the influence that coatings may have on substrate properties such as mechanical strength, resistance to creep and resistance to mechanical and thermal fatigue. Finally, it is stressed that proven performance for a given coating/substrate combination is no guarantee that no deleterious reaction will occur when the same coating is used with a different substrate alloy. Therefore, coating substitution requires requalification.     

Lifetime extension prediction of the rejuvenated first stage gas turbine blades

ABSTRACT

This paper presents a novel method for the first stage gas turbine blades lifetime extension based on degraded GTD-111 substrates hardness evolution after the rejuvenation heat treatments (RHTs) process. Moreover, the RHT impact on directionally solidified (DS) GTD-111 nickel-based suparalloys homogenisation and blades hardness properties were revealed and discussed. The measured hardness values indicate that turbine blades hardness progress is closely related to the nucleation of topologically close packed (TCP) phases, irregular γ’ (Ni₃, Al) precipitates growth, and Cr-rich M₂₃C₆ Coarsening at grain boundaries. Aspects like γ′/γ eutectic regions rafting, cooling holes oxidation attacks, and formation of cavities within γ-matrix were also identified and analysed.     

Metallographic Approach to Hirbine Blade Life Time Prediction

An economically optimized and safe utilization of turbine components is of vital importance to both manufacturers and operators of aero and industrial engines. Calculations of design life must take into consideration variations in several factors, such as material properties, manufacturing processes and operational conditions. Adding safety margins into the calculation results, with necessity, in a conservative figure.

The calculated life can be confirmed, or unconfirmed, by examination and testing of randomly selected components during maintenance and overhaul. Techniques for these examinations and for evaluation of the results, in terms of life predictions are not yet fully developed.

Methodology and practical results are presented, with comments on prospects and limitations. Medium strength, wrought superalloys to modern, highly alloyed cast blades are covered. Thelife limiting mechanism of the blades in the program is normally creep. Life time extension by rejuvenation has proven successful, both technically and economically. This paper will describe the practical application of this proven metallographic approach to life time prediction and extension.     

Failure assessment of the first stage high‐pressure turbine blades in an aero‐engine turbine

In this paper, microscopic analysis and fatigue experiment were conducted to detect the damage degree of turbine blades after a 600 h service and determine the reliable service time. First, several service blades were cut into slices on different cross sections and microscopically examined. It is found that the γ′ phase particles are slightly coarsened and the γ′ phase parameters change with temperature and stress distribution. Then, fatigue test was conducted on service blades simulating the real working condition. The test result shows that the blades after a 600 h service would serve for 3596 h more until failure during actual flight. The ultimate fracture is mainly caused by the interaction of fatigue, creep and oxidation. Besides, the γ′ phase parameters change obviously compared with service blade without fatigue test. It indicates that the γ′ phase parameters could be used to evaluate the microdamage of service blades, which has great significance for service reliability of the turbine machinery.     

Refurbishing procedures for blades of large stationary gas turbines

Abstract

The design life for blades of large stationary gas turbines is at least 100000 h in creep terms. Typical damage occurring within the lifetime of gas-turbine blading may be hot corrosion and/or erosion, foreign-object damage, tip rubbing, and cracking caused by low-cycle fatigue, thermocycle fatigue, high-cycle fatigue, and creep crack growth. When applying refurbishing procedures both technical and economic aspects should be considered. The repair procedures available are welding, brazing, plasma spraying, recoating, combined with special heat treatment cycles, and occasionally intermediate hot isostatic pressing. The limiting factors for repair procedures are the mechanical properties and the hot corrosion behaviour of the refurbished parts. Examples are given of results from metallographic and laboratory tests on refurbished Kraftwerk Union (KWU) test specimens and blades in relation to the material, design, and stressing of the components. Typical examples of the refurbished parts are also presented. Service experience with refurbished blades in stationary gas turbines is limited at present. Therefore, KWU can only consider these procedures in the light of the growing knowledge of refurbishing processes and the operating experience of refurbished parts as well as the economic aspects.

MST/102