Implications of engine's deterioration upon an aero-engine HP turbine blade's thermal fatigue life

International Journal of Fatigue 22 (2000) 147–160 www.elsevier.com/locate/ijfatigue

Implications of engine’s deterioration upon an aero-engine HP turbine blade’s thermal fatigue life M. Naeem a

a,*

, R. Singh b, D. Probert

b

Central Technical Development Unit (CTDU), PAF Base Faisal, Shahrah-e-Faisal, Karachi, Pakistan b School of Mechanical Engineering, Cranfield University, Bedfordshire, MK 43 OAL, UK Received 26 March 1999; received in revised form 25 August 1999; accepted 25 August 1999

Abstract Possessing a better knowledge of the impacts of engine deterioration upon an aircraft’s performance as well as its fuel and component life usage, helps the users make wiser management decisions and hence achieve improved engine utilization. For a military aircraft, using a computer performance simulation, the consequences of engine deterioration on a high pressure turbine blade’s thermal fatigue life are predicted.  2000 Elsevier Science Ltd. All rights reserved. Keywords: Engine deterioration; Equivalent full thermal cycle; Thermal fatigue; Turbine’s entry temperature

1. Background With the break-up of the Warsaw pact, there has been a rapid move towards reducing defence budgets of all the concerned countries. Politicians and the press have referred to this as a ‘peace dividend’; the governments re-allocating funds from defence to social programmes. The resulting financial cut-backs require that managers of military forces are having to concentrate on reducing the life cycle costs of their programmes to meet their new objectives. One way whereby this is being accomplished is through better engine parts usage. This has involved re-ascertaining how the expected lives of components have been prescribed and investigations into ways in which operating costs may be reduced, either

Abbreviations: CEFF Cooling effectiveness; EDI, EI, FI Engine deterioration index, erosion index and fouling index respectively; EFTC Equivalent full thermal cycle; ENG For whole engine; FOD Damage resulting from the presence of a foreign object in the gas stream of the engine; HCF, LCF High cycle or low cycle fatigue respectively; HP, LP High pressure or low pressure respectively; HPC, HPT High pressure compressor or turbine respectively; IECMS Inflight engine condition monitoring system; LPC, LPT Low pressure compressor or turbine respectively; PLA Power-lever angle; TET Turbine’s entry temperature, (K); TMF Thermal mechanical fatigue. * Corresponding author. Tel.: 0092-21-111444222 ext 2403; fax: 0092-21-9218323. E-mail address: [email protected] (M. Naeem)

through incurring less wear, decreased maintenance or extending the safe-life expectancies of components. All of these require accurate information concerning the condition of the engine and the stresses to which it is likely to be exposed. The effects of the reductions in military spending are being felt throughout the aerospace industry, resulting in increased competitiveness and a hesitancy to engage in high risk engine developments. Consequently contractors are trying to achieve enhancements of the useful lives of in-service (i.e. already in use) components, thereby achieving improved performance or safety. Regrettably, this tends to be expensive and may not be in the best interest of the individual user in terms of reducing total life cycle cost. Consequently, the onus is on the user to demand developments with high benefit-to-cost ratios. Much of this endeavor has been devoted to enhancing gas-path techniques, material-property improvements, and engine component usage analyses. In-service deterioration of any mechanical device, such as an aircraft’s gas-turbine engine, is inevitable. However, the extents to which such deteriorations adversely affect (i) the fuel and life usage, (ii) the aircraft’s operational performance as well as (iii) the overall life cycle costs, all remain to some extent esoteric. Hence this investigation was undertaken.

0142-1123/00/$ - see front matter  2000 Elsevier Science Ltd. All rights reserved. PII: S 0 1 4 2 - 1 1 2 3 ( 9 9 ) 0 0 1 0 5 - X

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Nomenclature b Fatigue strength exponent c Fatigue ductility exponent C Coefficient dependent on the material Cp Specific heat of gas at constant pressure (J kg⫺1 K ⫺1) E Elastic modulus (N m⫺2) High pressure spool speed, (rpm) N2 Nf Number of cycles before failure ensues Nf1, Nf2 Number of cycles to failure, for thermal cycle types 1 and 2 respectively P Larson–Miller parameter r Mean radius of rotation of the blade (m) tf Period to failure, (h) T Temperature, (K) T0, T1, T2 Maximum temperature of the metal blade for the reference thermal cycle or for thermal cycles 1 and 2 respectively, (K) Tblade Metal blade’s temperature, (K) Tcool Temperature of the cooling air prior to being influenced by the blades (i.e. HPC’s exit temperature), (K) Tgas Relative gas temperature, (K) ⌬e Strain range Fatigue ductility coefficient ef sf Fatigue strength coefficient w Angular velocity of the rotor

2. Analysis strategy Implications of engine deterioration upon the fuel usage and the aircraft mission’s operational effectiveness have already been investigated [1,2]. Because of the enormous effort required to carry out a comprehensive analysis of life expectancy, an investigation has been undertaken separately for creep, low cycle fatigue and thermal fatigue. Implications of engine deterioration upon the creep life and LCF life consumption of a HPT blade have already been investigated [3,4]. This paper discusses only the implications of engine deterioration upon thermal fatigue life of a HPT’s blade.

expired critical component and this is likely to lead to other parts of the engine being replaced prematurely. Another means of improving safety would be to design the turbine’s blades and discs very conservatively, i.e. return to the early design methodology of blade and disc operating at low stress levels. This design route would enable long and hence economic lives to be achieved. However, the weight penalty so incurred would reduce the payload of the aircraft and hence (i) the revenuegenerating capability of a civil aircraft, and (ii) the performance of a military aircraft.

4. Component degradation 3. Safety versus operating cost Deciding what should be chosen as an appropriate ‘working life’ of a critical component of a gas-turbine engine for an aircraft involves achieving an optimal balance between safety and profitable operation. The degree of safety achieved can be increased by using conservative criteria, which will translate into the components only being used for short service-lives and so possessing extremely low probabilities of failure. However, such short service-lives will result in more frequent and hence more costly programmes of replacement of the critical components. This will also lead to more frequent maintenance activities caused by the need to remove the life-

This is caused by the combined effect of the flight loads imposed, thermal distortions, erosion of airfoils, engine fouling due to internal deposits, in-service damage and abuse, the type of engine operation as well as duty cycle implemented and the maintenance practices employed for the engine [3]. Sallee [5] produced several performance deterioration models for the JT9D engine’s behaviour. These models show that the performance loss mechanisms associated with (i) the compressors cause reductions in both flow capacity and efficiency; and (ii) the turbines result in a flow capacity increase and an efficiency fall. The (i) and (ii) situations worsen upon increasing the number of flight cycles.

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5. Operating procedures to reduce engine deterioration Two factors that govern the rate of deterioration of an engine are the ways in which the engine is operated and maintained. Adherence to specified repair schedules, requiring frequent cleaning and reconditioning of components will reduce the rate at which performance deterioration occurs. The responsibility for the rate of deterioration must also be borne by the user, as the manner in which he / she operates the engine will have the greatest effect on the rate at which the engine will degenerate. Sallee [6] made the following recommendations in order to reduce the rate at which a gas turbine will degrade: 앫 After starting the engine, operate it at idle power for a minimum period of 5 min before accelerating. 앫 The initial acceleration from idling, for a repaired engine, should be achieved by a gradual incremental power increase. 앫 Unnecessarily large accelerations should be avoided: (i) Whenever possible: only a small acceleration or deceleration should be employed, e.g. a minimum of 60 s should be allowed before requiring the achievement of full power. (ii) When instantaneous decelerations are required, they should be accomplished as soon as possible after reaching high power settings. 앫 Each engine calibration should be undertaken for a decreasing power sequence, so that the engine will be relatively cool for shut down. 앫 Allow the engine to idle for 5 min prior to shutting down (in order to avoid abrupt large reductions in the temperatures of its components). Although these operations may be appropriate for some industrial gas-turbine applications, and for some transport aircraft, the majority of them are not feasible for an operational fighter aircraft. 6. Life-limiting failure modes of an aircraft gasturbine engines The durability of an aircraft’s gas-turbine engine is a function of its in-service resistances to its components’ failures, which mostly are dependent upon the operating environment of the engine. In order to predict the likely serviceable life of an engine component, it is essential to first understand its actual in-service usage. 6.1. Short-life failures These occur after a brief period of operation and are usually associated with the incorrect application of the

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design rules, a non-fulfilment to determine accurately the likely loads to be imposed on the component or a dereliction to manufacture the component to specification. Typical examples of such failures are: 앫 Excessive elastic deflections, causing components to jam or suffer excessive rubbing. 앫 Elastic instability leading to buckling. 앫 Plastic instability leading to necking. 앫 Gross plastic deformation leading to yielding. 앫 Fast fracture by an unstable crack propagation.

6.2. Non-localized damage This can be corrosion, erosion, FOD and/or uniform creep. The first three are typically functions of the external environment of the engine. Of these, only corrosion has an impact on LCF failure, as fatigue damage is typically exacerbated in a corrosive environment [7]. As these three failure modes are functions of external factors, then their exact prevalence can be difficult to predict. They are usually dealt with by periodic inspections and other maintenance actions, such as anti-corrosion washing. 6.3. Localized damage There are two significant localized damage mechanisms, namely creep and fatigue. 6.3.1. Creep The application of steady loads at high temperatures can cause cracks to nucleate and grow. This form of creep is macroscopically localized, unlike uniform creep. Creep is particularly prevalent in the turbine’s rotor and stator blades, although it can ensue in other components, e.g. turbine discs. The creep life tf has been described by the Larson–Miller equation: 103P logtf⫽ ⫺C T

(1)

where P is the Larson–Miller parameter, which is a function of the operating stress. As can be realized from this equation, a slight increase in the operating temperature T can lead to a large reduction in creep life. 6.3.2. Fatigue Fatigue failure is caused by the application of varying loads to a component, so producing fluctuating stresses. Aero engines are usually subjected to non-steady loads, which produce fluctuating stresses within components. If the fluctuating stresses are high enough, eventually they will cause failure, even though the maximum stress is less than the static strength of the material. A failure

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induced in this manner is called a fatigue failure. Low cycle fatigue is associated with relatively low numbers of high stress applications [4]. A fatigue is typically described as a LCF when the number of cycles to failure is fewer than 50 000. High cycle fatigue is associated with a high frequency of application of low stress loads. HCF failures are often caused by vibration, which cause high frequency low amplitude loads. Unlike LCF, HCF is characterized by spending the greater proportion of time in the crack initiation and short crack growth modes. 6.4. Thermal fatigue The most severe stresses that turbine blade encounter are those induced by extreme temperature gradients and rapid thermal transients. These thermally-induced stresses, combined with high mechanical loading, result in localized high transient strains and thereby produce thermal mechanical fatigue (TMF) cracking in the blade, especially if repeated often enough. The cracks tend to be initiated on the component’s surface (because of the highest temperatures occurring there) and subsequently propagate through the bulk of blade [8]. In advanced fighter aircraft engines, the components that come into direct contact with the high temperature gases (i.e. the turbine blades and nozzle guide-vanes, which are completely immersed in the hot gas stream, as well as the turbine discs, which are partially immersed) are highly susceptible to thermal mechanical fatigue. The high thrust-to-weight ratios, which are achieved primarily by increasing the combustor’s exit temperature, enhance the manoeuvrability of these aircraft, but this involves the turbomachinery components being cycled rapidly and more frequently. Consequently, the lives of high pressure turbine blades are dictated by a combination of cyclic thermal, centrifugal, and gas bending loads, or thermal mechanical fatigue (TMF). The F404 HP turbine blades (considered for investigation) incorporate complex internal cooling channels, and therefore thermal mechanical fatigue is also produced by the restriction of free expansion due to the local constraints induced by the complex cooling passages and the design performance requirements of thin leading and trailing edge profiles (especially as a result of the enhanced thermal gradients due to the high degrees of external heating and internal cooling). High pressure turbine alloys experience a wide variety of thermal and mechanical loading cycles depending upon the speed of the particular power transient and the part of the blade. In addition, the phasing of the thermal and mechanical cycles varies, the extremes being completely in-phase and completely out-of-phase. With inphase strain and temperature waveforms, the maximum load occurs at the maximum temperature, whereas with out-of-phase strain and temperature waveforms, the

maximum load occurs at minimum temperature [9]. The most deleterious case is the completely out-of-phase cycle due to the high mean stresses developed as a result of stress relaxation at the maximum cyclic temperature [10]. The damage caused by thermal mechanical fatigue, like creep, rises approximately exponentially as the temperature is increased [8]. Therefore, a small reduction in engine deterioration may have a dramatic effect on prolonging the component’s thermal fatigue life. The stages involved in calculating the transient thermal stresses are as follows [11]: 앫 The computation of the gas stream temperatures, pressures and velocities from the known engine characteristics. 앫 The deduction of the appropriate hot gas to metal components heat-transfer coefficients. 앫 The determination of transient temperatures, throughout the considered component, during each flight using evaluated gas stream temperatures and heattransfer coefficients. 앫 The calculation of the resulting thermal stress distribution profiles (for the various stages during the flight). In practice, this procedure is complicated and requires major computing power. This means that, for on-board real-time processing of thermal transient stresses in an aero-engine life usage monitoring system, a simplified thermo-mechanical model has to be developed. 7. Temperatures of the metal blades The temperature at any location in the considered blade is influenced by the temperatures of (i) the gas flow around the outside of the blade, and (ii) the coolant at the inlet to, and the outlet from, the blade. In developing the computer program for the determination of metal blade’s temperature distribution, the following factors were considered [8,12]: 앫 The HPT’s blades are cooled through multiple pass convection cooling and leading edge film cooling: the relatively cool air employed will increase the blade’s resistance to both creep and thermal fatigue. 앫 Cooling air is injected into the blades near their leading and trailing edges, i.e. the regions of the blade which are exposed to the greatest thermal assaults. 앫 Cooling air is bled from the last stage of the HPC. The combined convection/film cooling effectiveness has been defined [8,12] as CEFF⫽

(Tgas−Tmetal) (Tgas−Tcool)

(2)

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The cooling effectiveness of a rotor blade is dependent upon the total gas and coolant temperatures relative to those of the blades. If it is assumed that the axial velocity does not change across the rotor, and the tangential velocity is zero, the rotor’s relative gas temperature may be written as a function of the high pressure spool speed [12]: i.e.

冉 冊

冤 冥

2prN2 60 (wr) Tgas⫽TET⫹ ⫽TET⫹ 2Cp 2Cp

冋 册 2

2

(3)

It is assumed that the temperature of the coolant gas will be the same as the exit temperature of the gas leaving the HPC. It is not unreasonable to assume that the CEFF at each operating point remains almost invariant. Even though the HPT blades may be degraded, it is also assumed that their cooling effectiveness will not change. The metal temperature as a function of the gas and the coolant temperatures, and the CEFF (with a typical value of 60%) can be represented by the following simple equation [8]: Tblade⫽Tgas⫺CEFF(Tgas⫺Tcool)⫽Tgas

(4)

⫺0.60(Tgas⫺Tcool) It is envisaged that the temperatures calculated by this method will be overestimates, because no account is taken of the rate at which heat is absorbed by or released from the blades. As such the blade metal temperatures calculated will exceed the true values during engine accelerations but be lower during engine decelerations. However, for the present purpose this is acceptable because the aim of this study was to determine the percentage changes arising due to engine’s deterioration.

8. Assumptions, considerations and limitations A component does not have to fracture in order to fail in service: its performance can deteriorate to such an extent that it requires replacement. A failure due to component deterioration, is defined, by its performance decline in relation to that which occurred when it was new. For the purpose of this investigation, it was assumed that component is replaced at a deterioration level of 10% for individual components and at 6% for whole engine. Gas-turbine engine components are subjected to a complicated combination of both steady and cyclic, ther-

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mal fatigue and inertial stresses. These together with the effects of creep will eventually cause components to fail. In an effort to determine the safe-life limits, an understanding of the effects of these combined stresses is required—see Arvantis et al. [13] and Breitkopf et al. [14]. Arvantis et al. described a multi-axial life prediction system using a ductility–exhaustion method, that includes the evaluation of transient temperatures, corresponding elastic and inelastic strains, creep, and subsequently creep and fatigue lives. Breitkopf et al. used a method for determining the transient temperatures at different locations, deducing the thermal gradients from these temperatures and then superimposing the thermal stresses upon the fatigue stresses at these locations. These two methods illustrate that the analysis of complex stresses in gas-turbine engines is feasible but complicated, and requires access to data that are not readily available. As such, the majority of the literature tends to look at life limits based on more simplistic approaches. In this investigation, the combined effects of creep and fatigue are not considered. Rather each of the different failure mechanisms has been looked at separately. Implications of engine deterioration upon the creep life and LCF life consumption of a HPT blade have already been investigated [3,4]. This paper discusses the implications of engine deterioration upon the thermal fatigue life of a HPT’s blade. The effect of the environment also was not studied in this investigation. It is well known that the operating environment inside the hot gas path of a gas turbine is extremely hostile. The compound most commonly found is Na2SO4, which is formed in the combustion chambers of the gas turbine by the reaction of Na-containing salts, such as NaCl, which may be present in the air or fuel, with SO2 and SO3 created by the combustion of the sulphur-bearing fuel. In conventional operating environments, the conversion of NaCl to Na2SO4 occurs according to the following reaction [15]: 1 2NaCl⫹SO2⫹ (O2)⫹H2O→Na2SO4⫹2HCl 2

(5)

The Na2SO4 deposit is molten at the elevated temperatures occurring on the metal surfaces of the first stage vanes and rotor blades of the turbine. This modifies the oxidation conditions at the metal surfaces and so results in accelerated corrosion, because of the formation of sulphides at the interface between the metal and the oxide layer. Corrosion then penetrates along the grain boundaries into the metal, and also causes blistering of the protective oxide layer. Under a creep load, the stress field will promote the corrosion attack on the grain boundaries so leading to severe degradation [15]. This creep–corrosion interaction is a complex phenomenon beyond the scope of this investigation. Inclusion of heat transfer effects and signal lag times

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is very important for any analysis regarding thermal mechanical fatigue. However, the examination of the effects of an engine’s deterioration upon its life usage on a percentage change basis will be valid with or without the inclusion of these changes. As each engine will undergo the same transient series as the bench mark engine, it was assumed that approximately the same heat-transfer and signal time lag effects will be felt for each run. For these reasons, the effects of including heat transfer and signal lag should be considered as topics for future study. Life prediction for the complex thermal and mechanical loading situation is further complicated by the use of advanced anisotropic alloys and protective coating systems in these engines. Several approaches to the life prediction problem have been proposed [16–20]. However, these are frequently limited to specific test conditions, requiring a large data base or details of specific material characteristics/properties, which are not readily available in the open literature, for a successful generalized TMF life prediction system. A common postulate is that damage in a TMF cycle may be approximated as occurring at some ‘equivalent’ temperature (e.g. the maximum cyclic temperature) [10]. The method based on accumulation of equivalent full thermal cycles (EFTCs) was considered for the purpose of this investigation because: 1. It has actually been incorporated in the in-flight engine condition monitoring system (IECMS) algorithms used for the F404 on F-18 aircraft. 2. It predicts the relative severity of thermal fatigue. Therefore, although any inconsistency/discrepancy in the procedure will change the value of the life usage predicted, it will not influence the relative percentage differences brought about by the engine’s deteriorations (as considered in this investigation).

⌬e⫽b(T⫺T0)

where b is the coefficient of thermal expansion of the material of the blade, which is cycled between temperatures T and T0. The reference temperature T0 depends principally on the engine’s idling rpm at standard sealevel conditions. The strain range ⌬e is related to the number of cycles Nf completed before failure ensues [12] by: ⌬e sf ⫽ (2Nf)b⫹ef(2Nf)c 2 E

EFTCs provide an indication of the degree of thermal fatigue (which is the life-limiting failure mechanism for the F404 engine chosen to be investigated) experienced by the HPT’s blades. The EFTC is based on a prediction of the HPT metal blade’s leading edge temperature with the effects being measured relative to a reference cycle and summed to determine the amount of life consumed. Since thermal fatigue is the predominate life-limiting failure mechanism for the 64 HPT blades in the F404 engine, accurate monitoring of blade metal temperature is critical. Thermal fatigue is a condition whereby the strain cycles are predominantly caused by temperature cycling [8]. The total strain range is determined by:

(7)

where E is the elastic modulus of the blade’s material, and b and c for most materials equal ⫺0.12 and ⫺0.6 respectively. In Eq. (7), the first term results from elastic strain HCF and the second term from inelastic LCF. Generally, for aero-engines, the life of a HPT’s blade is long, e.g. for the F404 HPT blade, the life is 苲30 000 cycles. Therefore, it is assumed that the elastic term will dominate and hence, Eq. (7) may be re-written as: ⌬e sf ⫽ (2Nf)b 2 E

(8)

The strain range of any second thermal cycle, i.e. from T0 to T2 to T0, may be expressed [21] in terms of the first thermal cycle, i.e. from T0 to T1 to T0, as follows:

冋 册

⌬e2 T2−T0 Nf1 ⫽ ⫽ ⌬e1 T1−T0 Nf2

0.12

(9)

Nf1 is the relative severity of the second therNf2 mal cycle compared with that of the first thermal cycle. The thermal fatigue relative severity (R/S) can be expressed [21] in terms of the temperature differences as follows: The ratio

R/S⫽ 9. Equivalent full thermal cycle

(6)

冋 册

Nf1 T2−T0 ⫽ Nf2 T1−T0

8.333

(10)

The EFTC is based on calculating the severity of each of the cycles relative to a reference cycle and summing the total. Measurement of the thermal fatigue begins as the PLA passes through the idling setting (i.e. 60–70% of the design point LP spool speed for typical engines) and terminates when the cycle passes back through that point. For the present investigation, it has been assumed that the PLA is equal to and directly proportional to the LP spool speed. During this cycle (i.e. idling RPM to maximum RPM to idling RPM), the maximum value of the metal blade’s temperature is used in determining the thermal fatigue. The units of R/S are then accumulated to determine the thermal fatigue life usage for the HPT’s blades.

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10. Computer modelling and simulation For this investigation, the ‘Aircraft and Engine Performance—Simulation Computer Program’ (NaeemPAKa) evolved at Cranfield University has been used [12]. For the purpose of predicting the thermal fatigue relative severity for the HPT’s blades, the ‘Thermal Fatigue Relative Severity Prediction Model’ was developed. This computer program is written in the FORTRAN language and requires data for (i) the engine’s design and blade material (defined by the user through a separate file) and (ii) the time and speed history obtained from the ‘aircraft and engine performance simulation model’. For this investigation, a McDonnell Douglas F-18 aircraft, powered by two F404-GE-400 engines (referred to as F404s), has been assessed. The F404 is a low-bypass turbofan engine, consisting of six major modules and an accessories assembly, including (i) a three-stage fan driven by a single-stage LPT, and (ii) a seven-stage axial-flow compressor driven by a single-stage HPT. In practice, the set of basic performance parameters to be considered will depend upon the engine’s configuration. The two main engine parameters (namely the flow capacity and efficiency) are affected by the engine’s configuration and degradation. As yet, precise performance parameter changes due to typical faults in the engine’s components and their inter-relationships are not known accurately [22]. So, for this assessment, a ‘fouling index’, an ‘erosion index’ and an ‘engine deterioration index’ were used to describe the effects of changes in efficiency as well as the flow capacities of the compressor, turbine and whole engine respectively [4]. In general, unlike those for civil aircraft, the mission profiles for military aircraft can be relatively complex. However, for the purpose of the present analysis, a relatively simple aircraft mission profile has been assumed (see Fig. 1). Even a reasonably simple simulation can usually help in making ‘ballpark’ estimates. The basic methodology is to fly the aircraft through a complete mission profile with both engines (i) functioning properly and (ii) suffering prescribed identical degrees of deterioration.The chosen mission profile was used as the input to the computer simulation for the following sets of component and whole-engine conditions: 앫 clean engines, i.e. an EDI of zero for both engines; 앫 a FI of 1, 2, 3, …….. or 10% for each of either the LPC or HPC separately; 앫 an EI of 1, 2, 3, ……... or 10% for each of either the LPT or HPT separately; or 앫 an EDI of 1, 2, 3, …… or 6% for both engines. Subsequently the thermal fatigue relative severity prediction program is run using the predictions from the aircraft and engine performance simulation program as

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well as relevant material data as inputs. Thus the relative severity of the thermal fatigue that arises is predicted for each flight, so revealing the impacts of the engine’s deterioration on the thermal fatigue as a function of the aircraft’s flight path.

11. Discussion and analysis The prime aim of this study was to evaluate the impacts of various engine deteriorations upon the HPT blade’s thermal fatigue life. To carry out the required analysis, more than 150 aircraft and engine simulation runs were completed: each run took approximately 25– 30 min of computer time. As a result of the deterioration it has previously experienced, an aero-engine seeks a steady operating point which will differ from that for the same engine suffering no deterioration. This results in changes in the HPT’s rotational speed as well as the TET. Any rises in these two engine parameters increase the blade’s relative temperature and thereby result in greater thermal damage to turbine blades. The blade’s severity of thermal fatigue rises, with increasing engine deterioration for the engine as a whole, or for its compressor’s fouling or its turbine’s erosion (see Fig. 2). However, the extent of the rise at any given percentage level of deterioration depends on the type of deterioration. For example, with a 10% fouling index, the blade’s severity of thermal fatigue is 182 and 128% for LPC and HPC fouling respectively (expressed as a percentage of its value for the clean engine). This behaviour is because of the different reductions in the available thrust from the engines (at same TET), with increasing deteriorations of different types (see Table 1). The reduction in available thrust from the engines has a negative effect upon thermal fatigue life, because it requires the engines to run at higher TETs to match the thrust requirement in order to maintain the aircraft’s performance. Although the percentage reductions in available thrust for 10% deteriorations of either the LPC or the HPT performance are almost identical, i.e. 15% (see Table 1), the much higher blade’s severity of thermal fatigue with the LPC suffering fouling as compared with the HPT’s erosion (see Fig. 3) is because of the much higher HPT’s rotational speed with the fouled LPC (for the same developed thrust and at the same TET). A higher HPT’s rotational speed results in a higher blade’s temperature and thereby increased thermal fatigue of the blades. For example, at a TET of 1000 K, with a 10% LPC’s fouling or a 10% HPT’s erosion, the thrust developed is 4929 or 4935 kN respectively, whereas, the HPT’s rotational speed is 75 or 69.5% of the design point rotational speed respectively. It was also observed that the effects of component deterioration are not exactly additive (see Fig. 4). Usually, but by no means always, the majority of the

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Fig. 1.

Fig. 2.

Mission profile adopted (for this investigation) by the aircraft with clean engines.

Blade’s predicted severity of thermal fatigue for the stipulated deteriorations of the compressors, turbines or the whole engine.

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Table 1 Performance reduction as a result of the stipulated deterioration in the engines Component suffering deterioration

Deterioration level and type

Whole Engine LPC HPC LPT HPT

6% EDI 10% FI 10% FI 10% EI 10% EI

Reduction in the available thrust (expressed as a percentage of the available thrust for the engine, when clean) 35.7 14.9 17.0 9.5 14.8

fuel carried in the aircraft is consumed and time spent during cruise (e.g. 6355 s out of a total mission time of 8950 s in the assumed mission profile represents a major proportion of the total period to complete the mission). Therefore, it was considered appropriate to determine the effect of any change in the cruising Mach number and/or altitude would have on the blade’s thermal fatigue life. The blade’s severity of thermal fatigue (with the engines suffering a 6% deterioration) decreases slightly with increasing cruising altitude from 8000 to 11 000 m. After that, it rises initially at slow rate, but later at comparatively faster rate from 11 000 to 15 000 m (see the lower curve in Fig. 5). The shape of this curve is dictated by the overall effect of three main factors on

155

the TET’s variation. With respect to increasing cruising altitude of the aircraft, these are: (i) the reduction in the aircraft’s drag, (ii) the nearly linear decline in available thrust (for the same TET), and (iii) the fall in the atmospheric temperature (within the troposphere). The first factor is beneficial whereas the second has an adverse effect upon the thermal fatigue life, because respectively they require the engines to run at a lower and higher TETs, in order to satisfy the thrust requirement to keep the aircraft’s performance invariant. The favourable effect exceeds the adverse one up to the top of troposphere. However, the adverse effect exceeds the favourable one within stratosphere (i.e. above 11 000 m) because there the atmospheric temperature remains almost constant, but the aircraft’s thrust requirement becomes less because of decreasing atmospheric density with increasing altitude. The three factors mentioned above influence the TET’s variation mainly during the four flight segments shown in Table 2. Minimum values of the TET occur during flight segments (i) and (iii) at the cruising altitude of 11 000 m. This is because of the favourable effect of the reduced local atmospheric temperature at this altitude. Although the atmospheric temperature remains approximately invariant above 11000 metres altitude, reduction in the available thrust requires the engines to run at a higher TET in order to keep the aircraft’s performance constant. The overall impact of the TET vari-

Fig. 3. Severity of the blade’s thermal fatigue for the engine with a 10% FI for the LPC and HPC, and a 10% EI for the LPT and HPT separately (expressed as a percentage of the blade’s severity of thermal fatigue for a clean engine).

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Fig. 4. Blade’s predicted severity of thermal fatigue for engines with: a 6% FI for the LPC and HPC separately; a 6% EI for the LPT and HPT separately; and a 6% engine deterioration index (expressed as a percentage of the blade’s severity of thermal fatigue for a clean engine).

Fig. 5.

Blade’s predicted severity of thermal fatigue for the engines suffering a 6% deterioration when the A/C cruises at the stipulated altitude.

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157

Table 2 Maximum TET values and their differences for the clean and the 6% deteriorated engines during the stipulated flight segment of the mission profile Cruising Maximum TET and the difference in the values of the maximum TET (for the clean and the deteriorated engines) altitude (m) Cruise to a pre-set target (at 2100 Climbing from 8000 m to the Decelerating at the stipulated Descending from the stipulated km from home base) stipulated cruising altitude cruising altitude cruising altitude to 8000 m (i) (ii) (iii) (iv) Maximum TET (K)

8000 9000 10 000 11 000 12 000 13 000 14 000 15 000

0% EDI

6% EDI

1134 1119 1108 1103 1138 1193 1254 1328

1294 1277 1265 1258 1292 1351 1415 1498

Difference Maximum TET (K) in TETs

Difference Maximum TET (K) in TETs

Difference Maximum TET (K) in TETs

0% EDI

6% EDI

0% EDI

6% EDI

0% EDI

6% EDI

– 1409 1448 1490 1550 1597 1625 1633

– 1532 1553 1580 1654 1736 1822 1851

1120 1073 1059 1046 1065 1104 1162 1235

1272 1216 1200 1184 1211 1251 1313 1389

– 1057 1056 1054 1054 1054 1055 1055

– 1221 1220 1218 1218 1219 1220 1220

160 158 157 155 156 158 161 170

– 123 105 90 104 139 197 218

ation upon the blade’s thermal fatigue remains favourable from 8000 m until 11 000 m altitude. From 11 000 to 15 000 m altitude, an increasingly adverse effect ensues (i.e. the blade’s thermal fatigue increases, thereby shortening its life). The adverse effect of the engine’s deterioration upon the blade’s thermal fatigue life also rises with increasing cruising altitude (see the upper curve in Fig. 5). A greater difference in the TETs for the clean and the deteriorated engines (with the TET for the deteriorated engine being greater than that for the clean engine) means a larger adverse impact of the engine’s deterioration upon the blade’s thermal fatigue life (see also Table 2). For a given deterioration, the blade’s severity of the thermal fatigue increases (but with varying rate), with increasing cruising Mach number (see the lower curve in Fig. 6). This is due to increases of (i) the TET and (ii) the HPT’s rotational speed, in order to achieve higher cruising Mach numbers. For example, the maximum TET and the HPT’s rotational speed during cruising to a set target are 1693 K and 92% respectively, while the cruising Mach number is 1.5, as compared with 1261 K and 77.5% respectively while cruising at a Mach number of unity. The variations of these characteristics is mainly because of: (i) increasing of the aircraft’s drag; and (ii) decreasing available thrust (at the same TET) from the engines upon increasing the cruising Mach number. Less available thrust and a higher aircraft’s drag require the engines to run at higher TETs and HPT’s rotational speeds, thereby resulting in the blade’s greater severity of thermal fatigue. The reducing rate of rise in the relative severity of thermal fatigue from cruising Mach number 1.2 to 1.5 is because of the gradual reduction in the aircraft’s drag (and thereby the aircraft’s thrust

152 143 141 138 146 147 151 154

Difference in TETs

– 164 164 164 165 165 165 165

requirement) as a result of the gradual reduction in the aircraft’s drag coefficient [12]. The extent of the adverse impact of an engine’s deterioration on its blade’s severity of thermal fatigue changes with the changing aircraft’s cruising Mach number (see upper curve in Fig. 6). This is dictated by the overall effect of: (i) the difference in the TETs for the clean engine and after it had suffered deterioration, and (ii) the difference in the HPT’s rotational speeds for the clean and the deteriorated engines, at the stipulated aircraft’s cruising Mach number. A larger difference in the TETs and the HPT’s rotational speeds (with the TET and HPT’s rotational speed for the deteriorated engine being greater than that of clean engine) means a larger adverse impact of the engine deterioration upon the blade’s thermal fatigue-life. In this case, the variation in TET affect adversely, whereas the variation in the HPT’s rotational speed has a favourable impact upon the blade’s thermal fatigue-life (see Table 3). However, the overall effect of these two factors upon the blade’s thermal fatigue-life remains adverse throughout the range considered i.e. from a Mach number of unity to 1.5 number (see the upper curve in Fig. 6). As the ‘reheat-on’ flight segment influences the creep and LCF life so significantly [3,4], it was considered appropriate to see how the change in reheat-on time affects the blade’s thermal fatigue life. For this purpose, several simulation runs were carried out with different reheat times. The blade’s severity of thermal fatigue increases rapidly on switching on the reheat, but subsequently it increases more slowly with increasing reheat-on time (see lower curve in Fig. 7). The initial abrupt rise is because of the much higher TET and HPT’s rotational speed then attained compared with those of without reheat-on. The subsequent rises in the

158

Fig. 6. ber.

M. Naeem et al. / International Journal of Fatigue 22 (2000) 147–160

Blade’s predicted severity of thermal fatigue for the engines suffering a 6% deterioration when the A/C cruises at the stipulated Mach num-

Table 3 Difference in the values of the maximum TETs and the HPT’s rotational speeds (for the clean and the 6% deteriorated engine) Aircraft’s cruising Mach number

1.0 1.1 1.2 1.3 1.4 1.5

Difference in parameter value

TET for the clean engine minus the TET for the deteriorated engine (K)

HPT’s rotational speed for the clean engine minus the HPT’s rotational speed for the deteriorated engine (expressed as a percentage its value at the engine design point)

⫺158 ⫺147 ⫺127 ⫺124 ⫺136 ⫺136

3.6 5.3 6.6 7.0 7.0 7.2

TET and HPT’s rotational speed are much less with increasing reheat-on time (see Table 4). The adverse impact of the engine’s deterioration on the blade’s thermal fatigue decreases with increasing reheat-on time (see upper curve in Fig. 7). There are two factors responsible for this behaviour: (i) the difference in the increased TETs of the clean and deteriorated engines, and (ii) the difference in the increased HPT’s

rotational speeds for the clean and the deteriorated engines. The difference in the TETs remains almost same. However, the difference in the HPT’s rotational speeds decreases with increasing reheat time, thereby resulting in a lower blade’s severity of thermal fatigue (i.e. a lesser effect of the impact of engine deterioration upon the blade’s thermal fatigue life). The blade’s relative severity of thermal fatigue with a 6% engine deterioration decreases to 223% for a reheat-on time of 30 s as compared with that of 278% without reheat on (both expressed as percentages of their severities of thermal fatigue for a clean engine). LPC deterioration is the most serious in terms of its effects upon the blade’s thermal fatigue life. Therefore, in addition to the impacts of the LPC’s fouling, which results from combination of deteriorating flow capacity and decreasing efficiency, the impacts of these two parameters upon the blade’s severity of thermal fatigue were also analysed separately. For a LPC’s 5% flow-capacity deterioration, the blade’s severity of thermal fatigue increases more than linearly, with increasing efficiency deterioration (see upper curve in Fig. 8). On the other hand, for a LPC’s 5% efficiency–deterioration, the blade’s severity of thermal fatigue increases initially but almost immediately starts decreasing at a faster rate, but subsequently at a comparatively lower rate upon increasing the LPC’s flow capacity deterioration (see lower curve in Fig. 8).

M. Naeem et al. / International Journal of Fatigue 22 (2000) 147–160

Fig. 7.

159

Blade’s predicted severity of thermal fatigue for the engines suffering a 6% deterioration: effect of the stipulated reheat-on period.

Fig. 8. Blade’s predicted severity of thermal fatigue for the engines suffering a 5% LPC’s flow capacity and efficiency deterioration separately for the stipulated LPC’s flow capacity and efficiency deterioration respectively.

160

M. Naeem et al. / International Journal of Fatigue 22 (2000) 147–160

Table 4 The maximum TET and HPT’s rotational speed for an engine suffering a 6% detrioration (when the reheat is switched on for the sipulated period) Reheat-on time TET (K) (s)

HPT’s rotational speed (expressed as a percentage of its speed at the design point)

0 5 10 15 20 25 30

76.2 95.9 96.2 96.5 96.9 97.3 97.6

1253 1748 1761 1774 1792 1803 1816

12. Conclusions This investigation has quantified the impacts of the typical deteriorations, which an engine or its components suffer, upon the thermal fatigue-life of its high pressure turbine’s blade: LPC fouling had the most serious consequence. The individual thermal fatigue life shortening impacts of deterioration for all of the engine’s components for the HPT’s blades can not be summed or superimposed. The effects of aircraft’s cruising Mach number and altitude as well as engine’s reheat-on time, on the blade’s thermal fatigue-life were quite significant.

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