HAYNES ® 244 ® alloy was chiefly developed to address the need for high-strength, low coefficient of thermal expansion (CTE) alloys for seal rings and cases in advanced gas turbine engines. In addition to these attributes, adequate resistance to low-cycle fatigue (LCF) due to cyclic thermal and mechanical loading during service is critical for such applications. The isothermal LCF performance of commercially produced 0.5” (12.5 mm) thick, fully heat treated plate products of 244 alloy was evaluated by means of axial strain-controlled (R = −1) LCF tests covering total strain ranges up to 1.25 % (without dwells), at temperatures ranging from 800–1400°F (427–760°C). In addition, the comparative LCF performance of Waspaloy, a well-established alloy for turbine cases, was evaluated under selected, nominally identical test conditions. S-N curves were constructed and fitted by the Coffin-Manson equation, allowing the delineation of regimes controlled by the elastic and plastic response of the material. Fracture surfaces were examined in the scanning electron microscope to identify fatigue crack initiation sites and crack propagation modes. Differences between the alloys are discussed in terms of tensile strength and cyclic hardening/softening behavior. Implications for fatigue performance of these alloys under cyclic thermal loading conditions are discussed as well.

The symptoms of the propagation of shaft cracks in rotating machines, which can be brought out of the vibration data collected by condition monitoring systems, are often not very evident, especially when the crack depth is not very high. In fact, shaft cracks commonly propagate rather far from the cross-sections where the vibration probes have been mounted. At these measurement planes, the sensitivity of the machine dynamic behavior to the effects caused by a shaft crack can be fairly low. Therefore, it is very important to improve the capabilities of diagnostic techniques aimed at the early detection of shaft cracks. In this paper, the dynamic behavior of the cracked shaft of a HP-IP steam turbine is shown. The results of investigations on the influence of the periodic and non-periodic breathing of the crack on the rotor response, measured in different machine thermal states and at unusual cross-sections of the shaft, are shown and discussed.

Additive Manufacturing (AM) applications are poised to rapidly expand in aero-engines, providing opportunities to manufacture complex geometries which are unobtainable through traditional manufacturing techniques. The powder based AM process, selective laser melting (SLM), results in susceptibility to defects such as porosity, lack of fusion, and unmelted powder. These random defects often serve as fatigue crack initiation sites, thus influencing fatigue crack initiation and growth behavior. Consequently, the influence of random defects on high cycle fatigue performance for the AM titanium compressor disk is investigated in this study. First, the relationships between the defect area and the fatigue performances involving fatigue limit and the crack propagation threshold are quantified. Second, the effect of defect probability distribution is examined. Finally, the durability analyses on SLM titanium compressor disks in Ti-6A1-4V alloy are performed based on the fatigue life prediction model for defect-containing AM material. Then sensitivity analysis is conducted to determine the critical defect size influencing the fatigue limit and fatigue life.

Fatigue life assessment of turbine rotating components is usually performed taking into account the presence of potential defects and their propagation, together with an accurate description of the stress/strain fields acting at components critical regions. In this work, such an approach was applied on a radial outflow turbine: the component investigated in this paper is the tie-rod. An axisymmetric FEM model of the whole turbine was implemented to evaluate the stress field acting at the root of tie-rod thread, the component most critical region. Three different preload procedures were investigated, in order to evaluate the effects of plastic strains on fatigue cycles. It was found that an initial overtightening, performed before machine assembly, induces a larger plastic zone at the root of the thread, which lowers the peak stress and lessens the stress ratio of the fatigue cycle. Stress gradients for the three different preloading conditions were the starting point of a crack propagation model, which calculated crack advancement as a function of service time. The model, which considered crack closure and stress ratio effects, allowed structural integrity assessment and showed that the initial overtightening had also the effect of crack growth rates reduction.

Fatigue crack growth in plastic regions is one of the main factors for life assessment of components subjected to plastic strains. Residual life can be evaluated considering a crack growing from component first load cycle, considering an appropriate crack growth model. The aim of this work is to discuss the effects of temperature on crack propagation during low cycle fatigue. In particular, a series of experiments at high temperature (up to T = 700°C) were performed on a Ni-based superalloy, to determine crack growth rates in the LCF regime. During tests interruptions, crack growth was measured using the plastic replica technique. Experimental results were examined in terms of crack growth rates and a model based on elastic-plastic J-Integral. Results showed that at high temperature there is a significant increase of crack growth rates respect to models usually valid at room temperature.

During rapid engine throttling operations, turbine airfoils can experience very rapid heating and cooling, particularly at take-off conditions. These rapid transient events lead to the generation of high thermal gradients and non-uniform stress distributions through the thermal barrier coating, environmental barrier/bond coating and substrate. This in turn can lead to coating delamination, overheat of the substrate materials, creep and thermo-mechanical fatigue of the part. We present the process and computer modeling methodology for a physics-based prediction of deformation, damage, crack propagation and local failure modes in coated turbine airfoils and other parts operating at hot section turbine environment conditions as a function of engine operational regimes, with a particular emphasis on rapid transient events. The overall goal is to predict the effects and severity of the cooling and heating thermal rates on transient thermal mechanical fatigue life of coated hot parts (turbine airfoils, blade outer air seals and combustor liners). The computational analysis incorporates time-accurate, coupled aerothermodynamics with non-linear deformation thermal-structural finite element modeling and fracture mechanics modeling for high rate thermal transient events. Thermal barrier coating thermal failure and spallation are introduced by the use of interface fracture toughness and interface property evolution as well as dissipated energy rate. The spallation model allows estimations of the part remaining life as a function of the heating/cooling rates. Applicability of the developed model is verified using experimental coupons and calibrated against burner rig test data for high flux thermal cycles. Our results show a decrease in TBC spall life due to high rate transient events.

High velocity oxygen fuel (HVOF) sprayed stellite coating is often adopted on contact surfaces of the shrouds and stubs of titanium turbine blades, in order to increase fretting fatigue strength and reduce fretting wear. To confirm the effectiveness of the sprayed stellite coating, we conducted friction-type fretting tests originally developed to simulate the load condition of shrouds and stubs, vibratory force was carried only by friction force. We also investigated the fretting fatigue mechanism of coated materials by observing non-propagating cracks in the tested specimens and analyzed crack propagation behavior at the interface using fracture mechanics. The fretting fatigue tests confirmed that the sprayed stellite coating could double the fretting fatigue strength of Ti-6A1-4V. The fretting fatigue mechanism of Ti-6A1-4V with a sprayed stellite coating was revealed as follows: Final fracture was led by a newly initiated micro-crack at the interface as cracks initiated on the coating surface did not continuously propagate into the substrate but mainly propagated along the interface. The micro-crack propagation at the substrate depends on the length of interface crack. The longer interface crack delays micro-crack propagation to the substrate because the lines of force are less concentrated at the crack tip. Hence, the interface crack makes the fretting fatigue strength of coated Ti-6A1-4V higher than that of uncoated specimens. This finding was supported by the cross-sectional observations of non-propagating cracks in the tested specimens, and by FE analysis which revealed that Δ K of the substrate crack initiated at the interface decreases with an increase in delaminated crack length. Ti-6A1-4V with a sprayed stellite coating offers better wear resistance than uncoated Ti-6A1-4V because the coating’s friction coefficient is lower than that of Ti-6Al-4V, although the wear rates of coated and uncoated materials against the consumption energy per cycle were almost the same.

This paper discusses the methodology to calculate high cycle fatigue (HCF) crack propagation life of gas turbine bolts and compares two dimensional (2D) HCF crack propagation life to three dimensional (3D) HCF crack propagation life. Gas turbine bolts when exposed to fatigue loading are prone to crack initiation and propagation (structural failure) during operation. In such cases cracks mostly are initiated by low cycle fatigue (LCF) and propagated by HCF. Therefore in current illustration the authors have evaluated crack propagation primarily initiated by low cycle fatigue and propagated by high cycle fatigue. 2D and 3D fracture methodology approaches had been used for analytical evaluation. The authors conclude on the efficacy of both the methods based on the data from the field. The coupling joint bolts located in the engine mid-section, which are used to join compressor rotor with turbine rotor are being considered for crack evaluation studies. The coupling bolts located in mid-section are primarily loaded by high axial bolt pre-loads needed to keep the joint intact, as well as loaded in bending due to rotor gravity sag. The crack propagation life is evaluated and validated with field data using cracked bolt specimen from the field.

CSP plants using supercritical CO 2 (sCO 2 ) power cycle can potentially achieve high thermal conversion efficiency at low capital cost due to compact turbomachinery and other components. An sCO 2 expander and improved heat exchanger is expected to provide a major stepping stone for achieving CSP power at $0.06/kW-hr LCOE, energy conversion efficiency > 50%, and total power block cost < $1,200/kW installed. However the life limiting mechanisms of these turbomachines in high pressure, high temperature sCO 2 environment are not well understood. To understand the effect of high pressures, high temperatures and sCO 2 chemical kinetics on crack initiation, crack propagation and low cycle fatigue (LCF) life of these turbomachines, a novel experimental setup is developed. Advanced microstructure and spectroscopic analyses are conducted that shed light on some key differences between various Ni base alloys in terms of oxidation morphology, chemical species diffusion and trapping, the formation of protective corrosion resistant layers and changes in surface properties. An experimental technique for low cycle fatigue experiments in high pressure, high temperature supercritical CO 2 environment is developed. The test setup allows for pressurized LCF testing of alloys being considered for MW scale sCO 2 turbine development. Results show that the LCF life remains the same (within the scatter band) irrespective of the location of crack initiation site whether at the OD (non shot-peened bars in air and sCO 2 ), or at the ID (shot peened bars). Total fatigue life, for all conditions, lie within the normal variation in LCF results (± 2X life variation). No significant LCF life debit is observed in IN718 by sCO 2 at 550 °C, 0.7% max strain, 20 cpm. Similar conclusion is reached during 0.6% max strain tests. The effect of sCO 2 is found not to be significantly more damaging than air at these strain levels. However, the results can be different for lower % max strains due to longer exposure times involved, resulting from larger number of cycles to failure. Similarly at higher temperatures and/or longer hold-times, sCO 2 environment may be more aggressive, resulting in lower total fatigue life.

The combined high and low cycle fatigue (CCF) test on full scale turbine blade in the laboratory is an important method to evaluate the life. In fact, the low cycle fatigue which is usually caused by the centrifugal force can be confirmed easily. While, the high cycle fatigue which is usually caused by the vibration and aerodynamic force is often hard to determine. So the previous scholar has proposed the contrast method to determine the high cycle load in the field. This method utilizes the new and used blades to determine the high cycle within certain limits. While it can’t be applied effectively in the whole life range with the low cycle-high cycle-ultra high cycle fatigue theory raised. So this paper put forward the modified contrast method to realize the optimization. Firstly, the CCF tests are carried out on the turbine blade systematically. Then, the CCF damage properties, including the crack propagation, the fracture morphology and the dynamic characteristic are analyzed. Lastly, the new modified contrast method is proposed with the new coordinate axes, new fitting criterions and amend method. Through comparisons we conclude that: the new method is slightly complicated, but the evaluate precision has significantly increased. So it could be used to deal with data for CCF tests on full scale turbine blade in the future.

Spallation of thermal barrier coatings subjected to thermal fatigue occurs through cracking in or close to the metal–ceramic interface. To better match an experimentally observed damage progression curve with initially slow crack growth followed by rapid crack growth, a multitude of crack paths were modeled in a finite element analysis. Comparisons with experimental data enabled the most likely crack path to be identified. It was shown that the most likely failure type for the studied TBC system was crack initiation from pre-existing defects in the thermal barrier coating which propagated into the interface, leading to spallation. The results were used to fit a crack propagation model.

Ceramic turbines could substantially increase operating temperatures of turbomachinery without the need of blade cooling, leading to higher conversion efficiency and power density. However, the inherent brittleness and low tensile strength of ceramic materials limits the use of hub-based ceramic turbines. This paper presents a novel inside-out turbine architecture, permitting the use of monolithic ceramic blades. The proposed architecture uses a carbon-polymer composite rim which converts the rotor radial loads to tangential hoop stress. The blades mainly support compressive loads, minimizing tensile stresses within the blade and thus crack propagation. This allows the use of low tensile strength ceramics which cannot be used in standard hub-based turbines. The rotor hub is comprised of two radially flexible C-shaped hubs, which have sufficient compliance to follow radial displacement of the heavily loaded composite rim. The feasibility of the proposed inside-out ceramic turbine is demonstrated by addressing the four key challenges of the architecture using proof-of-concept prototypes, namely: (1) rotor dynamics of the flexible hub; (2) thermal viability of the composite rim (3) local tensile stress in the blades, and (4) thermal shock in the ceramic blades in transient mode. Experimental validation with an alumina blade confirms that this architecture supports the use of low tensile strength, brittle ceramic blades.

The demand for higher plant cycling operation and reduced life-cycle costs are the main drivers for the design and assessment of turbine components today. Heavy cyclic loading increases the potential of fully utilizing the fatigue capabilities of the material which might lead to crack initiation and subsequent crack propagation. Fracture mechanics methods and evaluation concepts are widely applied to assess the integrity of components with defects or crack-like findings. The realistic modelling of the failure mechanism plays a key role for the accurate prediction of crack sizes at failure state. A basic treatment of material toughness typically leads to conservative assessments for components with sufficient ductility. A standard approach to describe material behavior with high ductility is to use the start of stable crack extension as a dimensioning parameter for the analysis. By definition a critical condition for a component is reached when the crack driving force is equal to the characteristic material parameter. On the other hand, advanced analysis methods allow determination of the instability point (ductile tearing analysis). This paper will discuss two cases for practical analysis from steam turbine design showing clear advantages for service application by using advanced analysis methods.

U.S. Department of Energy (DOE) has recently sponsored research programs to develop Megawatt scale supercritical CO2 (sCO2) turbine for use in concentrated solar power (CSP) and fossil based applications. From rotordynamics perspective, the extremely high power densities associated with this sCO2 turbine (similar to those of rocket turbo pump) and the long shaft of the turbine require mid-span support from bearings to achieve the high rotational speeds needed to achieve desired turbine efficiency targets. The mid-span bearings, if oil lubricated, need two sets of seals to isolate the working fluid from lubricating oil and represent a combined parasitic load of as much as 5% of the power plant output. These parasitic loads can be avoided by the use of hybrid gas bearing (HGB) which uses turbine working fluid itself as the lubricant for generating hydrostatic + hydrodynamic pressure and hence can operate in a high temperature, high pressure sCO2 environment. A successful and reliable implementation of HGB is expected to provide significant efficiency improvement in the sCO2 turbine for modular plants in 10MW scale. However, the behavior of hybrid gas bearing with supercritical CO2 as the working fluid has never been studied before. Hence a careful investigation of its hydrodynamic and rotordynamic characteristics with perturbations associated with sCO2 fluid and sCO2 turbine is warranted. In this paper the use of HGB to bring the critical speed ratio of highly power dense 10 MW sCO2 turbine into a manageable regime is investigated. Further, the performance characteristics of HGB in high pressure, high temperature sCO2 are investigated with a coupled fluid-structure interaction model. Differences between load capacity and stiffness characteristics for traditional operation in air versus operation in highly dense and viscous sCO2 are highlighted. Using Bayesian probabilistic models, insights into key design parameters needed to optimize HGB performance are obtained. A 3D fracture mechanics and crack propagation model is also developed to predict life of HGB under typical sCO2 turbine mission cycle for CSP application.

Heavy duty gas turbine blades and vanes are operated at high temperatures and high stresses, condition where several damage mechanisms can simultaneously be present. For example creep, fatigue and oxidation play an important role in the propagation of existing cracks. Crack growth models are employed for assessment criteria, interpretation of the field feedback and non-conformities management and they are required to be as accurate as possible when predicting crack propagation under the combined effect of all the three phenomena. In this work, a Linear Elastic Fracture Mechanics (LEFM) model based on isothermal experimental tests and validated by Thermo-Mechanical-Fatigue Crack Growth tests (TMFCG), is employed to predict crack propagation of a cast Ni-base superalloy used in gas turbine blades and vanes. When calculating the individual propagation fractions of creep and fatigue crack growth, the model accounts for the instantaneous stress state and temperature in transient regime (i.e. a complete cycle of start-up, base-load and shut-down). The loss of γ’- precipitates at the crack tip due to surface oxidation is interpreted as environmental damage fraction. A complete workflow for the systematic use of the approach, comprising an in-house software, has been defined and developed. Stress intensity factors used for LEFM calculations are determined either using tabulated weight functions or with the aid of Finite Element Analysis (FEA). This flexible approach is consistent with the industrial need of a given fracture mechanics calculation, which might require different levels of accuracy and resources/time consumption case by case. The software identifies the fraction of propagation caused by oxidation, creep crack growth or fatigue crack growth. This allows checking the physical realism of the results by comparing to metallographic analysis of fracture surfaces from broken TMFCG test specimen and/or real component damage information from field. Besides, this feature can be helpful to support the engineer in residual life evaluation under damage tolerant approach because it allows the identification of the type of operational regime that minimizes crack propagation. The software also allows the execution of sensitivity analyses via Monte-Carlo calculations, identifying for a given component and operational condition the more relevant calculation inputs. This feature also quantitatively supports the engineers in the identification of the most appropriate safety margins.

Finite element simulations based on an interface cohesive zone model (CZM) have been developed to mimic the interfacial cracking behavior between the α-Al 2 O 3 thermally grown oxide (TGO) and the aluminum rich Pt–Al metallic bond coat (BC) during cooling from high temperature to ambient temperature. A two dimensional half-periodic sinusoidal geometry corresponding to interface undulation is modelled. The effects of TGO thickness and interface asperity on the stress distribution and the cracking behavior are examined by parametric studies. The simulation results show that cracking behavior due to residual stress and interface asperity during cooling process leads to stress redistribution around the rough interface. The TGO thickness has strong influence on the maximum tensile stress of TGO and the interfacial crack development. For the sinusoidal asperities, there exist a critical amplitude above which interfacial cracking is energetically favored. For any specific TGO thickness, crack initiation is dominated by the amplitude while crack propagation is restricted to the combine actions of the wavelength and the amplitude of the sinusoidal asperity.

The failure of vital components is not uncommon in the gas turbine industry. In the event excessive degradation occurs within a component, e.g. extensive cracking in a turbine blade or vane, solutions exist to either repair or replace defective parts. Such parts are readily accessible and mostly exchangeable in the field, limiting the amount of outage time and assessment required for defective parts. When more critical components exhibit extreme wear or cracking, e.g. a crack in a rotor disk, repairs typically necessitate a complete rotor destack and refurbishment or have the potential to require the replacement of individual disks. In extreme cases, defects found in rotor disks can be known to retire an entire compressor or turbine rotor. The OEM solution of replacing disks puts a substantial cost on the customer, thus providing an incentive for characterization and advanced analyses to determine the residual life in critical rotating components. Considered an advanced analysis, linear elastic fracture mechanics (LEFM) provides the theory and fundamental structure to conduct crack growth analyses in components that exhibit nominally elastic behavior. Successful implementation of LEFM requires extensive characterization of the material, engine operating boundary conditions, and high fidelity finite element models. Upon the detection of a flaw, whether an internal or external indication, the results from finite element analyses can be used to derive the crack tip stress field and subsequent crack tip driving parameters. These parameters are then utilized in a comprehensive crack propagation model, calibrated to temperature- and load-dependent material data, to determine the number of cycles to unstable propagation. As a result, the remaining life of a component with a given indication is readily obtained, enabling our engineering team to provide a thorough life assessment of critical rotating components. An overview of the linear elastic fracture mechanics crack growth analyses conducted is presented, with a special emphasis on compressor and turbine disks.

The framework for characterizing and predicting failure when simple crack propagation control predominates is reasonably well understood. Since, in many ceramic systems, crack propagation appears to be the primary failure mode much of this paper is devoted to a description of the crack propagation process. Other failure modes, such as crack initiation and complex crack propagation control, are, unfortunately, in a rather rudimentary state of development; however, some discussion of situations where these might occur is presented.

During the operation of the gas turbine engine, the crack initiation sites have frequently been found at the slot bottom in turbine disks. Taking the cracks at above mentioned position of the existing turbine disk as an example and considering the feature that the cracks usually initiate early and grow slowly, this paper searches for the initiative cause and the propagation law of such cracks. The results obtained can be employed as an important basis for damage tolerance or fracture control for engine design.

Incorporation of the Turbine Engine Structural Integrity Program (ENSIP) has provided a more organized and disciplined approach to the durability related aspects of the engine demonstrator program. As a result, engine capabilities are more easily related and the number of durability problems is reduced. This is accomplished by increased contractor/customer interface in the design analysis, material characterization and verification testing of the engine programs. The demonstrator program has been affected by the ENSIP requirements, particularly the Damage Tolerance aspects. The design analysis is more complex while the configurations tend to be simpler with fewer notched locations. Materials are developed for improved crack propagation life while component and engine testing is expanded to verify this capability. The ENSIP specification effects increases in analysis, materials development, verification testing and durability.