Patent Grant
10890083
The present disclosure concerns a method of controlling turbine tip clearance and a corresponding turbine tip clearance system.
In a gas turbine engine there are various rotary components, including compressor stages and turbine stages. Each of these stages comprises a hub and an array of blades extending radially away from the hub. The blades each have a root portion at the hub, an aerofoil portion having pressure and suction sides, and a tip at the distal end to the hub. The array of blades is surrounded by a casing which is static. A clearance is required so that neither the blade tips nor the casing is eroded by tip rub from contact between them. Nevertheless, it is beneficial that the clearance between the tips of the blades and the casing is minimised so that all the working fluid, for example air, passes over the aerofoil portions and does useful work rather than leaking over the tips.
An initial tip clearance can be designed. However, during use of a gas turbine engine the blades and casing expand and contract at different rates to each other, due to thermal and centrifugal loading, and so the tip clearance changes. Typically the tip clearance of the turbines at take-off of an aircraft powered by one or more gas turbine engines is large, because the engine is relatively cool, and the tip clearance reduces at subsequent flight conditions, such as climb and cruise.
Turbine tip clearance systems are known to modulate the tip clearance of turbine rotors to reduce the gap to a safe minimum to improve turbine efficiency without threatening tip rub. Active and passive turbine tip clearance systems are known and may comprise, for example, cool air blown towards the casing to arrest its growth or shrink it radially and thereby reduce the radial distance between the blade tips and the casing.
According to a first aspect of the invention there is provided a method of controlling turbine tip clearance comprising steps to:
Advantageously the method enables the turbine speed and turbine temperature to be managed so that the available margins for both speed and temperature run out at the same time. In particular the turbine tip clearance can be used to trade between the turbine temperature and the turbine speed.
The steps of measuring the turbine speed, measuring the turbine temperature, measuring the parameters, and determining the speed and temperature limits may be performed in any order.
The turbine speed may be measured once per flight or may be measured once every n flights. The turbine speed may comprise the peak turbine speed. The turbine temperature may be measured once per flight or may be measured once every n flights. The turbine temperature may comprise the peak turbine temperature. The engine deterioration which reduces the available margin for turbine speed and turbine temperature occurs over a relatively long period of operation so it is not necessary to measure turbine speed or temperature frequently. The measured turbine speed and turbine temperature may be used to calculate the offset to the target tip clearance for the next flight or flights. Advantageously the calculation therefore has a different iteration rate to the main turbine tip clearance control loop so the dynamics of each loop do not adversely interact. Advantageously taking fewer measurements reduces the processing power required. Advantageously the sensors for measuring the turbine speed and/or turbine temperature may be extant sensors. The measurements may also be used for other purposes within the engine control and/or monitoring.
The parameters may comprise engine power level. The parameters may comprise flight condition. The parameters may comprise altitude and/or throttle position.
The step to calculate target tip clearance may comprise using a proportional-integral algorithm. Advantageously this provides closed loop control of the tip clearance and therefore minimises the requirement for accurate data tuning on initialisation or subsequently.
The step to calculate target tip clearance may comprise an input that is the difference between the current measured tip clearance and the calculated target tip clearance. This is an offset.
The step to measure turbine temperature may comprise measuring one or more engine parameters and deriving turbine temperature therefrom. The engine parameters may include compressor delivery temperature; compressor delivery pressure; fuel flow rate; nozzle guide vane area. Advantageously the method is also applicable to turbines where the temperature is too high to measure directly, or for which it is more convenient to measure different engine parameters and to derive the turbine temperature.
According to a second aspect of the invention there is a turbine tip clearance system comprising:
Advantageously the system enables better management of turbine speed and temperature margins. Advantageously the system also enables the turbine efficiency to maintained within the limit constraints.
According to a third aspect of the invention there is a turbine assembly comprising the turbine tip clearance system of the second aspect.
The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein.
Embodiments will now be described by way of example only, with reference to the Figures, in which:
With reference to
The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by suitable interconnecting shaft 23, 24, 25.
Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
For the present disclosure the gas turbine engine 10 is assumed to power an aircraft, either singly or with one or more identical engines.
Each of the fan 13, intermediate pressure compressor 14, high pressure compressor 15, high pressure turbine 17, intermediate pressure turbine 18 and low pressure turbine 19 comprises one or more rotor stages. A schematic illustration of a rotor stage 28 is shown in
In use of the gas turbine engine 10, working fluid (air) does work on the rotor blades 32 as it passes substantially axially through the engine 10. Working fluid that passes over the blade tips 34 through the clearance 38 does no useful work and therefore reduces the efficiency of the engine 10 and increases fuel consumption. However, the clearance 38 is necessary to prevent the blade tips 34 rubbing against the rotor stage casing 36 which causes damage to one or both components. Tip rub is a transient effect because the rub erodes the blade tip 34 or casing surface which results in the clearance 38 being increased and therefore the engine efficiency reducing.
Additionally the clearance 38 is not constant throughout use of the gas turbine engine 10. Taking the example of a gas turbine engine 10 used to power an aircraft, the rotor stage 28 components grow and shrink in response to centrifugal forces and temperature changes resulting from different engine operating conditions. Thus when the engine 10 is cold, before use, the rotor blades 32 have a defined radial length and the rotor stage casing 36 has a defined diameter and is annular. The components each grow or shrink by different amounts and with a different time constant governing the speed at which the growth or shrinkage occurs. The growth due to centrifugal forces is substantially instantaneous.
When the engine 10 is switched on it begins to heat up and the rotor hub 30 and blades 32 begin to rotate which causes all the rotating components to grow radially. Due to the rotation of the rotor blades 32 and their relatively small mass the rotor blades 32 tend to grow radially very quickly, substantially instantly, by a small amount. The hub 30 grows radially outwardly by a relatively large amount, for example three times as much as the rotor blades 32, with a long time constant, for example 100 seconds. The rotor stage casing 36 which is relatively massive and does not rotate grows by a relatively large amount, for example three times as much as the rotor blades 32, but with a long time constant, for example 50 seconds. The segment assembly grows by a small amount, for example a third of the growth experienced by the rotor blades 32, with a moderate time constant, for example 15 seconds.
During engine acceleration the rotor stage casing 36 and hub 30 grow radially outwardly and the rotor blades 32 elongate radially. The segment assembly may grow radially inwardly. The net effect is that the clearance 38 increases during engine acceleration phases of the flight such as ramp up and the start of take-off because the growth of the casing 36 is larger and quicker than the growth of the other components. Similarly, the clearance 38 decreases during engine deceleration phases because the casing 36 contracts more quickly than the hub 30. There is a settling period after an engine acceleration or deceleration during which the clearance 38 may fluctuate before settling to a steady-state clearance 38.
It is known to apply active or passive tip clearance control arrangements to reduce the variation of clearance 38. For example cool air can be selectively delivered to passages in the rotor stage casing 36 to cool the rotor stage casing 36 and thereby reduce the diameter or retard the growth of the diameter. Alternatively the segment assembly radially inside the rotor stage casing 36 can be moved mechanically to change the clearance 38.
Turbine temperature TGT is the temperature of the gas that enters the low pressure turbine 19. The temperature TET of the high pressure turbine 17 exhibits a similar profile but is lagged in time. TET may be calculated from other engine parameters including a temperature measured in a cooler part of the engine 10, for example the temperature T30 (sometimes called T3) and pressure P30 (sometimes called P3) of the compressor delivery air where it is delivered from the high pressure compressor 15 into the combustor 16, the fuel flow Wf and the area of the nozzle guide vanes for that turbine 17. Although turbine temperature TGT is used herein it will be understood that the temperature at the entry to the high pressure or intermediate pressure turbines 17, 18 may be used instead.
A turbine case cooling system 80 is generally indicated in
A turbine tip clearance system 80 is shown in
The turbine tip clearance system 80 may deliver cool air to shrink the turbine casing 36 in some engine operating phases. It may also or alternatively be arranged to deliver relatively hot air to the turbine casing 36 in order to cause it to expand more rapidly. This is particularly important during rapid but transient acceleration phases of engine operation where the rotor blades 32 and hub 30 of the high pressure turbine 17 may grow more quickly than the turbine casing 36. The turbine tip clearance valve 90 is actuated to open or close in order to deliver control air to the casing 36 or to prevent control air from reaching the casing 36. The turbine tip clearance valve 90 may be bi-stable so that it is either open or closed. Alternatively it may have more than one open position or be continuously variable to direct controllable amounts of air to the turbine casing 36. The turbine case cooling valve 90 may be arranged to always deliver a small flow. Thus there is no closed position but instead is a minimum flow position and one or more positions providing more cooling flow. There may be more than one turbine tip clearance valve 90, which preferably are mutually controlled to act in concert.
The turbine tip clearance system 80 includes a processor 86 which performs steps of the method 54, which is described below. The processor 86 may be common with the controller 84, or may be coupled thereto. A memory may be a part of the processor 86 or may be operatively coupled to the processor 86. The processor 86 is coupled to one or more sensors 88 from which the measurements used in the method 54 are obtained. Thus there is one or more sensor 88 which measures the rotational speed NH of the high pressure shaft 23. There is also one or more sensor 88 which measures the turbine temperature TGT. There may also be one or more sensors 88 which measure the compressor delivery temperature T30, compressor delivery pressure P30, fuel flow Wf and nozzle guide vane area A. These are engine parameters from which turbine temperature TGT may be calculated. There is also one or more sensor 88 which measures parameters indicative of flight conditions and engine power level.
The turbine tip clearance system 80 may be part of a turbine assembly. The turbine assembly may comprise one or more rotor stages 28 each having a hub 30 and an array of blades 32 radiating therefrom, a turbine casing 36 radially outside the tips 34 of the rotor blades 32, and the turbine tip clearance system 80.
In the graph shown in
The high pressure shaft 23 couples the high pressure turbine 17 and the high pressure compressor 15. They rotate at a common high pressure speed, NH, which is quicker than the intermediate pressure shaft 24 or the low pressure shaft 25. This high pressure speed NH increases rapidly from ground idle to take-off of the aircraft. As can be seen in
Similarly the low pressure turbine temperature, TGT, increases rapidly from ground idle to take-off of the aircraft. It also exhibits an overshoot, albeit small in the example illustrated, which straddles the time marked take-off and then reduces to a steady value.
At the beginning of take-off the turbine tip clearance valve 90 is at its minimum flow position so minimal, or no, control air is delivered to the turbine casing 36. This is because the cold build tip clearance is sufficient that no tip rub will occur during the start or ground idle phases of engine operation. At a predefined time during take-off the valve 90 is opened so that cool air is delivered to the turbine casing 36, which can be seen by the line 40 on
The actual tip clearance 38 is initially the cold build clearance. As the engine 10 accelerates from ground idle the high pressure turbine 17 and turbine casing 36 grow radially due to thermal and centrifugal loading. In the absence of clearance control the tip clearance 38 initially decreases during acceleration (due to centrifugal growth) and then increases because the casing 36 grows thermally more quickly than the hub 30 and blades 32. The turbine tip clearance valve 90 may be opened when a threshold is reached, for example when the turbine speed NH has increased to a predetermined level. Alternatively the turbine case cooling valve 90 may be subject to closed loop control such that the actual tip clearance, whether measured or calculated, is one of the inputs to the control of the turbine case cooling valve 90. The control may include a predictive element which uses models of the component growths to predict how the tip clearance 38 will change in a defined time horizon and then control the turbine case cooling valve 90 to alter that predicted behaviour.
The actual tip clearance 42 may generally be considered to be the minimum tip clearance measured to the tips 34 of the rotor blades 32, which therefore accommodates any non-concentricity between the rotor blade tips and the turbine casing 36. Due to the cooling air flow provided by the opening of the turbine case cooling valve 90, line 40, the growth of the casing 36 is arrested or slowed. The hub 30 generally has a slower time constant for its growth and so it continues to grow radially so that the blade tips 34 are moved radially outwardly. As a consequence the tip clearance reduces, as can be seen from line 42 in
Once the turbine tip clearance valve 90 has been opened the actual tip clearance 42 continues to reduce but the rate of decrease slows. At the point where the actual tip clearance 42 begins to level off the turbine tip clearance valve 90 may be partially and/or progressively closed to reduce the amount of control air that it delivers to the high pressure turbine 17. This is shown in
By tightly controlling the actual tip clearance 42 using the turbine tip clearance valve 90 the high pressure turbine 17 is efficient during take-off. Known turbine tip clearance systems seek to optimise the efficiency n of the high pressure turbine 17 by controlling the actual tip clearance 42. However, controlling the actual tip clearance 42 to optimise the turbine efficiency n causes the relatively large overshoot of the turbine speed NH, and reduction of the turbine temperature TGT overshoot, during take-off. This is illustrated in
As a consequence of the elevated turbine speed NH the components are exposed to higher centrifugal loading which can have a detrimental effect on rotor disc life. The high pressure compressor 15 which is coupled to the high pressure turbine 17 via the high pressure shaft 23 also rotates more quickly and therefore the speed ratio between the high pressure compressor 15 and the intermediate pressure compressor 14, and/or between the high pressure compressor 15 and the combustor 16, may be off-design and therefore less efficient than is achievable at lower speed. In addition, certification requirements set limits on the turbine temperature TGT and turbine speed NH which are permissible.
The method herein seeks to sacrifice a small amount of turbine efficiency q in favour of maintaining both turbine speed NH and turbine temperature TGT below acceptable limits. Advantageously the method may be applied only where the engine 10 is operating close to either or both of the limits during take-off. Where the operation is away from both limits the method may be omitted.
The relationship between turbine speed NH and turbine temperature TGT is fixed by the physical design of the engine 10. The high pressure turbine 17 operates at a fixed NH/√TGT ratio. However, during take-off transient effects such as changes in the air system flows, compressor and turbine capacities, and the efficiencies of the compressors and turbines can shift the matching between the turbine speed NH and the turbine temperature TGT. The lines 48 show a set of the turbine speed NH and turbine temperate TGT matching options during take-off. The arrow 50 indicates the progression between the lines 48 which results in increasing turbine efficiency η. Thus line 48a represents a less efficient operation of the high pressure turbine 17 than line 48c because it has a lower turbine speed NH for a given turbine temperature TGT, or because it operates at a higher turbine temperature TGT for a given turbine speed NH.
Control area 52 is the part of the engine operation where the high pressure turbine 17 is close to either or both of the NH limit 44 and TGT limit 46. Optionally the method is applied only when it is detected that the high pressure turbine 17 is operating within control area 52. Alternatively, the method may be applied during all take-off operations but may include suitable filters so that the effect is only apparent close to the limits 44, 46. There is also a reduction in compressor delivery temperature T30 when the turbine efficiency q is reduced by reducing the turbine speed NH. Thus it may be beneficial to operate the method during all take-offs.
Over time the high pressure turbine 17 operation moves from relatively low turbine speed NH and turbine temperature TGT during take-off, towards the origin of
The method seeks to sacrifice some turbine efficiency η in order to move the high pressure turbine 17 operation onto the line 48b so that both limits 44, 46 are reached approximately simultaneously. Since the turbine speed NH overshoot is greater than the turbine temperature TGT overshoot, when tightly controlling actual tip clearance 42, the turbine speed NH may be reduced at the expense of a small increase in turbine temperature TGT to move the relationship from line 48c to line 48b.
The method uses the tip clearance control system 80 to actively manage the turbine speed NH and turbine temperature TGT as well as the turbine efficiency q. Thus the method can choose to control to a non-optimal turbine efficiency η in order to reduce the turbine speed NH. This may increase the turbine temperature TGT.
The method 54 also requires measurement of the turbine speed NH, box 58. This may be obtained using any known method. For example, there may be a phonic wheel mounted to the high pressure shaft 23 with a sensor mounted to static structure close to it in order to record the time of arrival of each tooth and thence derive the rotational speed. Alternatively a sensor may be mounted radially outside the turbine blade tips 34 or compressor blade tips to record the times of arrival and thence derive the rotational speed. Alternatively the turbine speed NH may be measured via a gearbox mounted to be driven by, or to drive, the high pressure shaft 23. In this case a speed sensor or secondary PMA winding is provided in the gearbox to provide a speed measurement which is multiplied by the gear ratio between the gearbox and the high pressure shaft 23. Where the gearbox is mounted to a different shaft 24, 25 of the gas turbine engine 10 direct measurement of the shaft speed may be necessary.
The method 54 requires a measurement of the turbine temperature TGT, box 60. This may be obtained by direct measurement using a thermocouple or other suitable temperature probe. Alternatively, particularly where the operating temperature of the turbine is too high to be directly measured, the turbine temperature TGT may be derived from other engine parameters. Thus there may be an optional precursor step to measure, or otherwise obtain, the relevant engine parameters and to derive or calculate the turbine temperature therefrom.
The TGT limit 46 is obtained at box 64. Similarly the NH limit 44 is obtained at box 66. The limits 44, 46 may be stored in memory and retrieved therefrom.
The measured and/or calculated turbine speed NH and turbine temperature TGT may first be used to determine whether the high pressure turbine 17 is operating within the control area 52. The rest of the method 54 may be applied only if the operation is within the control area 52. Alternatively the method 54 may be applied for all take-off operations.
The turbine speed NH and turbine temperature TGT from boxes 58 and 60, the parameters from box 56 and the limits from boxes 64, 66 are all inputs to a function to calculate and optimise the turbine tip clearance; box 68. Finally the calculated and optimised turbine tip clearance is passed to the turbine tip clearance system 90 to control the turbine tip clearance apparatus, box 70.
Alternatively the offset, box 72, may be determined by inputting the turbine speed NH and turbine temperature TGT to a proportional-integral (PI) controller. An example of this type of controller is shown in
The offset, box 72, and target tip clearance, box 57, are added together in a summer 74 and the resulting clearance is outputted to a control algorithm 76 as shown in
It may be beneficial to calculate the offset, box 72, only once per flight, or once per n flights where n>1 since deterioration affecting the offset changes over a number of operating cycles, and to use the calculated offset in the next flight. Thus the offset could be calculated using the peak turbine speed NH_peak and peak turbine temperature TGTpeak; that is the maximum values in a flight or cycle, or in a set of flights or cycles. An example of this type of controller is shown in
In
Whichever function is used to calculate the turbine tip clearance 38 also seeks to optimise the turbine efficiency q within the constraints of the NH and TGT limits. Thus the function may first seek a turbine tip clearance 38 which maximises the turbine efficiency q and then adjusts the turbine tip clearance 38 to ensure both turbine speed NH and turbine temperature TGT remain below their respective limits 44, 46 with the consequent reduction in turbine efficiency q. Alternatively the function may iteratively seek values of turbine tip clearance 38 that meets the constraints of the limits 44, 46 and the requirement to have good turbine efficiency q. In either case the aim is to optimise the tip clearance 38 for the best turbine efficiency q possible whilst meeting the TGT limit 46 and the NH limit 44 constraints. It is also the aim to reach the TGT limit 46 and the NH limit 44 approximately simultaneously. This can also be considered to be running out of margin for turbine temperature TGT and turbine speed NH simultaneously.
In other applications it may be desirable to seek to optimise the operational cost within the TGT and NH limits 46, 44. Such an optimisation may take account of the effect on life of the turbine components by increasing or decreasing the turbine temperature TGT and turbine speed NH relative to each other. Beneficially this may result in lower operating cost and/or longer intervals between engine maintenance actions because part life is improved.
The calculated and optimised tip clearance, box 68, is provided to control the turbine tip clearance apparatus, box 70 of
Although a gas turbine engine 10 for powering an aircraft has been used to describe the features of the invention, the method and apparatus are applicable in other contexts. For example, the method can be applied to a marine gas turbine engine 10. The method can also be used in an industrial gas turbine engine 10, particularly one which is configured to supply peak power and therefore accelerates rapidly from stationary with the consequent rapid thermal and centrifugal growths affecting the turbine tip clearances 38.
Although the operation has been described in relation to acceleration during take-off of an aircraft, it is also applicable in other contexts. In particular the method may be applied whenever there is a large transient of the engine operation which requires the tip clearance 38 to be controlled in order to maintain efficient operation and to prevent tip rub. Such transient may be causes by a large power off-take, particularly an unscheduled power off-take. For example, in a military aircraft a large power off-take may be demanded to power a weapon mounted to the aircraft. In a marine gas turbine engine large power off-takes may also be demanded.
It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.
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