GB2445565A

GB2445565A - Gas turbine engine having a plurality of modules comprising a combustor and transition duct

Info

Publication number: GB2445565A
Application number: GB0618904A
Authority: GB
Inventors: Senior Peter; Nigel Wilbraham
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2006-09-26
Filing date: 2006-09-26
Publication date: 2008-07-16
Also published as: GB0618904D0; WO2008037554A2; WO2008037554A3

Abstract

A gas turbine engine includes a number of high-temperature modules, each of which includes a combustor unit 20 and a transition duct 22. The modules are secured to the engine by means of a displacement mechanism 28, which enables rapid and easy access to the module and adjacent inner parts 18, 26 of the engine. The displacement mechanism is preferably a pivot arrangement 28, by means of which the module is rotatably attached to the engine frame. Clamps 33, 34 may be used to further secure the module to the engine casing. Preferably, the module also includes stator blades 24 which are attached to the end of the transition duct 22. Optionally, the displacement mechanism may comprise lugs engaging tabs (see figures 7a, 7b), or a spring. This arrangement enables quick and easy access to the combustor, transition duct, stator blades and rotor during repair. Flexible fuel lines 32 may connect the module to the fuel manifolds.

Description

IMPROVEMENTS [N OR RELAT[NG TO GAS TURBINE ENGINES The invention

relates to a gas turbine engine comprising a plurality of hightemperature modules.

The hottest components of a gas turbine engine include the last stator blading of the compressor, the combustor and the first stator blading and first rotor blading of the turbine. These components are subject to a gradual reduction in performance and integrity, due to the combined effects of elevated temperature and stress at the high pressures found in these parts of the engine. Unfortunately these components are usually the most difficult to remove for inspection., replacement, refurbishment or repair.

The time required to replace these parts significantly affects the availability of the engine for operation and represents a large proportion of the loss of utility to the customer. It is sometimes necessary to dismantle an engine on-site, in order to gain access to these components. Such dismantling can involve extra fixtures in order to stabilise the remaining pressure casing, as sections of the engine are removed. These fixtures are necessary to avoid movement in the engine structure, which can cause damage to the remaining rotor and blading. In those engines having a disc-rotor construction, it may be necessary to completely dismantle the rotor in order to gain access to the rotor blading. Re-assembly of the rotor will then require that the rotor be rehalanced, which adds to the maintenance costs.

Where engines have a welded-shaft type rotor, the same access usually requires that the upper casings be completely removed and that a heavy hoist be used to remove the rotor. The blading in the bottom half-casing may then be loosened and removed.

Where casings are removed, it is necessary to remove fuel manifolds and auxiliary equipment, in order to gain access to the casings, and special lifting arrangements must be made in order to cope with the weight of the equipment being lifted.

All of these procedures involve a substantial dismantling of the engine, which takes it out of service for a significant period. There is also the drawback that, when an engine is dismantled on-site, an error could easily be made in re-assembling the engine.

Furthermore, while the maintenance work is being carried out, it is possible for foreign bodies (large or small) to enter the engine, which can cause damage or even total destruction of the engine, if such bodies are not removed prior to re-commissioning. In addition, the space required on-site for carrying out such dismantling and for storing the necessary handling equipment is of concern for some customers, especially on offshore platforms, where space and weight are at a premium. It would therefore clearly be a great advantage if tooling and lifting gear could be eliminated altogether.

In some circumstances it may be deemed necessary to return an engine to base.

This adds transport time to the time during which the engine is out of commission, increasing maintenance costs still further.

The disadvantages involved in dismantling all or part of an engine, as outlined above, has led to some inspections being carried out by boroscope. However, this has the drawback that access to some major areas of the engine is restricted and also the visual inspection afforded by such a technique is not as discriminating as a direct inspection by the human eye. This increases the risk that a critical flaw, which could prejudice the integrity of an engine, could be missed.

The invention provides a gas turbine engine comprising a plurality of high-temperature modules, each module comprising a combustion unit for combusting a mix of a fuel and an oxidant, and a transition duct for conveying the combustion gases produced by the combustion unit to the rotor blading of the engine, wherein each module is secured to the engine by means of a displacement mechanism enabling rapid, easy access to the module and adjacent parts of the engine.

The end of the transition duct may be arcuate in shape, and an opening, which is provided in the engine for insertion or removal of at least parts of the module, may be configured so as to allow insertion or removal of the transition duct.

Each module may advantageously further comprise stator blading disposed at the end of the transition duct for receiving the combustion gases from the transition duct, the module directing the gases to the rotor blading. The stator blading may also be arcuate in shape, the opening being also configured so as to allow insertion or removal of the stator blading.

The opening may be rectangular or oval in shape.

The displacement mechanism preferably includes a pivot mechanism, which allows said at least parts of the module to be selectively inserted into or removed from the engine through the opening. The pivot mechanism may comprise a single pivot attaching the module to the engine. Alternatively, the pivot mechanism may comprise a first pivot attached to the engine, a second pivot attached to the module and at least one intermediate link between the first and second pivots.

The displacement mechanism may include at least two lugs attached to the engine at opposite ends of the opening and a corresponding number of tabs attached to the module, whereby the module can be temporarily held adjacent the opening prior to selective securing of the module to the engine or removal of the module from the engine.

The displacement mechanism may comprise a spring, one end of which is attached to the engine and the other end of which is attached to the module, whereby the module can be temporarily held adjacent the opening prior to securing of the module to the engine or removal of the module from the engine.

A clamp may be provided at a part of each module, which, when the module is seated against the engine, lies on an opposite side of the opening to the pivot mechanism., thereby to retain the module in its seated position. A further clamp may be provided at, or close to, the pivot mechanism, thereby to force the module against the engine at the location of the further clamp. At least one clamp may be provided in contact with the module, thereby to force the module against the engine.

The clamp or clamps may be a band clamp, which co-operates with all of the modules simultaneously.

Each module may comprise a flange, which is bolted to the engine casing.

A seal may be provided at an interface between a distal end of the stator blading and a part of the engine housing proximate to a first rotor blading of the engine. The seal may be secured to the distal end of the stator blading, being thereby removable as part of the module.

The seal may be a honeycomb seal. Such a seal may include among its cells solid stop elements, which are configured such as to prevent the cells from being crushed in a sealing state of the honeycomb seal.

The modules may be connected to fuel manifolds, which are located on a pivot side of the modules. The modules may advantageously be connected to the fuel manifolds by way of flexible fuel lines, such that the fuel lines do not have to be disconnected when said at least parts of the modules are removed from the engine.

The pivot mechanism may lie on a side of the module, which is located nearest the first rotor blading.

Embodiments of the invention will now be described, by way of example only, with the aid of the drawings, of which: Fig. 1 is a simplified axial sectional view of an embodiment of a gas turbine engine in accordance with the invention; Figs 2A and 2B are axial and radial sectional views of the embodiment of Fig. I, employing a squirrel-cage engine frame; Figs 3A, 3B and 3C are snapshots of a combustor-unit used in the embodiment of Fig. I before and during maintenance; Fig. 4A is a variant of the embodiment of Fig. 1 employing a honeycomb seal; Fig. 4B is an inset showing a detail of Fig. 4A, while Fig. 4C is a view of the honeycomb seal along an axial direction; Fig. 5 is a view of the combustor-unit illustrating the best positions for a pivot arrangement forming part of an embodiment of the invention; Fig. 6 is a view of the combustor-unit employing a variant pivot arrangement; and Figs 7A and 7B depict an alternative displacement mechanism to the pivot arrangement.

An embodiment of the invention is illustrated in Fig. 1. Fig. 1 shows a simplified sectional view of a gas turbine engine, consisting of three main sections: a compressor section 10, a combustion section 12 and a turbine section 14. The compressor section supplies air under pressure to the combustion section, where a mixture of the compressed air and incoming fuel is ignited, producing combustion products, which drive the turbine rotor 16. During this process compressed air is guided by a series of blading stages, of which the last stage is identified as item 18, to the combustion section 12. The combustion products from the burning of the fuel/air mixture in the combustor 20 are guided along a transition duct 22 through stator blading 24, thereby to impact on the first rotor blading 26 of the turbine. This and subsequent stator/rotor rows exert the necessary torque on the rotor shaft, turning the rotor.

The operations just described are, in fact, standard operations to be found in most gas turbine engines. The difference with the present invention, however, is that, whereas in conventional engines the combustor 20, transition duct 22 and stator blading 24 are fixed items, in the sense that, for example, the combustor is bolted down to the engine frame at the upstream end and the transition duct 22 and stator blading 24 are likewise attached to the engine frame at the downstream end, in the embodiment of the invention being described, the combustor 20, transition duct 22 and stator blading 24 all form one integral unit (combustor unit or "module"), which is displaceably secured to the engine by a pivot arrangement 28. Since these components are, as mentioned earlier, components which are especially exposed to high temperatures, the combustor unit may be termed a "high-temperature module". Thanks to this design, when it is desired to inspect critical components such as the compressor stator blading, the combustor and the turbine stator and rotor blading, all that the maintenance personnel have to do is unlock the combustor units and swing them out of the engine, leaving access through the combustor-unit openings to the afore-mentioned components. This greatly reduces the down-time of the engine compared with the known maintenance methods, and significantly reduces maintenance costs. Furthermore, in the conventional design it is necessary to provide for a dismountable joint between the combustor and the transition duct, because these are mounted to the engine frame at different points and differential movement takes place due to thermal and mechanical transients in the operating cycie of the engine. In the proposed embodiment, the transition duct 22 can be made integral with the combustor 20, avoiding the additional cost of the joint and eliminating a potential leakage path, thereby improving engine performance. Moreover, the actual movement of such joints in a conventional design leads to frictional wear, so that the proposed design will also improve the life of the combustion system.

The inclusion of the turbine first stator blading 24 along with the transition duct has the advantage that this allows this blading to be securely attached to the end of the transition duct, so that leakages and other undesirable heat-transfer disturbances between the transition duct and first stator blading are avoided. Thus, the performance of the engine and its life can be further optiniised. This arrangement also enables maintenance personnel to readily remove the first rotor blading of the turbine, assuming that this blading is secured to the rotor disc by a suitable blade-locking mechanism.

The combustors 20 are connected to fuel manifolds 30 via fuel lines 32. These lines may be inflexible, in which case it will be necessary to uncouple the lines from the combustors before the latter are pivoted out, or the lines may be flexible, in which case no such uncoupling will be necessary. Clearly, the latter option is to be preferred, since it makes it possible to gain even faster access to the inside of the engine, thereby reducing maintenance costs still further.

As is the case in conventional engines, the combustor units are secured to the engine through respective seal flanges (mentioned later on in connection with Fig. 7), which provide a gas-tight fit of the combustor units to the engine frame. In addition., the embodiment of the invention employs clamps in order to force the combustor units down onto their seals, thereby completing the sealing action, in the embodiment shown, these clamps comprise a first band clamp 33 and a second band clamp 34. The first band clamp 33 is placed over a point of the combustor units, which lies opposite to the pivot arrangements 28. Since this clamp is in the form of a band, it may be used to secure all of the combustor units at once against the engine frame, by being run circumferentially around the engine, over the just-described point of the combustor units, and tightened. In order to ensure that the combustor units are adequately sealed against the engine frame at the pivot location also, it is arranged for the pivot arrangement to have a certain amount of play and for the combustor units to be likewise urged against the engine frame through the seal at the pivot location by tightening a second band clamp 34 over the combustor-unit pivot arrangements 28. Thus, by applying and tightening the two band clamps 32, 34, the combustor units can be secured in sealed fashion to the engine. Since all combustors are released at once, with an operation (i.e. releasing the clamp) which is significantly faster even than unbolting a single combustor, very significant savings in downtime can be achieved. However, even if conventional bolting means, instead of clamps, were used to secure a seal, the pivot arrangement would still have advantages in terms of easier handling and inspection.

The combustors 20 may be cannular in design, which would normally require the use of a circular opening in the engine frame to accommodate the combustor.

However, the shape of the end of the transition duct, and indeed of the turbine stator blading 24 attached thereto, is conventionally arcuate, being often referred to as a "smile". The dimensions of such a "smile" are normally too great to allow removal of the transition duct or the turbine stator blading through such a circular combustor opening, which would normally only be large enough to allow the combustor (so-called "can") to pass through. Consequently, in the embodiment shown the opening is made non-circular, being in practice either oval or rectangular in order to allow removal of these "smile"-shaped components.

Figs 2A and 213 illustrate one way of configuring the combustor-unit opening.

The engine is constructed as a so-called "squirrel-cage" design, being divided circumferentially into a number of individual faces, one for each combustor. In the example shown, there are six faces for six combustors. The faces are divided into two sets of three, with flanges 42 on the upper and lower casings being provided to join the two halves together. The openings 40 for the combustors are, in this case, rectangular, though -as mentioned above -they could just as well be oval. Where a cannular-type of combustor is used, the combustor is mounted to a flange member, which is dimensioned to be slightly bigger than the opening, so that adequate sealing can take place. In a variant of this design, the flanges 42 are dispensed with, the engine then being of a unitary construction.

Figs 3A, 3B and 3C show the process of temporarily displacing the combustor unit out of its normal position sealed against the engine to a maintenance position, which allows access to the inner parts of the engine. Fig. 3A shows the combustor unit in its sealed position, with the band clamps in place and tightened. This is the normal operating position of the combustor unit. Figs 3B and 3C show two possible scenarios when maintenance work is to be carried out. The scenario of Fig. 3B is one in which flexible fuel lines are used. The two band clamps are in that case untightened and removed and the combustor units simply swung out of their sealed position so that the combustor itself and its associated transition duct and stator blading are removed from the engine entirely, leaving space for maintenance personnel to inspect the critical components inside the engine. While the combustor unit is thus swung out, the stator blading 24, transition duct 22 and combustor 20 can also be inspected, and even dismantled, if necessary. Since the fuel lines are flexible, they simply bend as necessary while the combustor unit is being pivoted out. Consequently, there is no need to uncouple them from the combustor.

The alternative situation is shown in Fig. 3C. In Fig. 3C the fuel lines are inflexible pipes and therefore are shown uncoupled, in order to allow the combustor unit to be swung out of the engine.

In the embodiment being described, the fuel manifolds 30 are shown as being on the side of the combustor facing the pivot arrangement 28. This is advantageous where flexible fuel lines are employed, since it allows these lines to be shorter, thereby saving on material costs. Were the pivot arrangement to be on the opposite side, where the clamp 33 is presently shown, the fuel lines would have to be long enough to allow them to extend further while the combustor unit was being swung open. This would therefore require longer fuel lines, increasing the material costs.

In addition to the seal already mentioned between the combustor flange and the engine housing, a further seal is required between the turbine first stator blading 24 and the area of the first rotor blading 26. One realisation of such a further seal is illustrated in Figs 4A-4C. This seal is in two parts: an outer part 50, which seals the end of the transition duct canying the first stator blading 24 to the engine housing; and an inner part 52, which seals the transition duct to the end of the rotor disc 54 -see Fig. 4A and enlarged representation of the sealing region shown in Fig. 4B.

Both parts of this seal are preferably a so-called honeycomb seal, consisting of a large number of cells disposed radially and circumferentially adjacent to each other, as shown in Fig. 4C. (Further details of a honeycomb seal can be found in, for example, United States patent application US 2005/0063816, published on 24th March 2005.) These cells extend axially referred to the longitudinal axis of the rotor and the seal follows a circumferentially arcuate shape, corresponding to the afore-mentioned "smile". The cells are hollow and the cell walls are resilient in an axial direction. In practice, because the cells are hollow, there is sufficient back-pressure set up by vortex-like flows inside the cells to greatly reduce the amount of leakage which occurs.

Included among the cells are solid elements ("stop" elements) 56 which are provided to prevent the cells from being crushed when the combustor arrangement is in its sealed position. In one realisation these stop elements have very little resilience and stand slightly proud of the hollow cells walls, so as to prevent any mechanical force on the honeycomb from occurring at all. However, as an alternative, the stop elements can be made to have some resilience, so that they accept most of the loading and allow the hollow cell walls to flex a small amount elastically (a recoverable deflection). This provides an even tighter seal.

The lower seal part 52 has a honeycomb structure like the upper seal part 50, but without the stop elements. This is because, in operation, this lower seal part does not actually contact the moving rotor disc, otherwise unacceptable power loss and wear would take place. Either the lower seal part can be dimensioned to include from the start a small axial clearance with respect to the rotor disc, or it can be made somewhat larger axially by the application of an abrasive coating. During rotation of the rotor, this coating then erodes away some of the seal part, leaving only the slightest clearance between the seal part and the rotor disc.

It was earlier stated that it was advantageous to have the pivot arrangement 28 on the same side as the fuel manifolds 30. It is also advantageous to have the pivot arrangement on the turbine side, rather than the compressor side, since the pivot will than be nearer to the interface between the turbine first stator blading 24 and the turbine first rotor blading 26. Since this area normally represents the area of greatest temperature, there will be a temperature differential between this area and other areas further upstream., i.e. in the direction of the compressor. Consequently, as was suggested earlier, slight movements (especially axially) in the rotor components and engine housing at the afore-mentioned interface area could give rise to significant relative movements between the first rotor blading 26 and the first turbine stator blading 24, which could prejudice the integrity of the seal at this interface area. When the pivot arrangement is on the turbine side, however, there is less distance between it and this interface area and therefore temperature differentials have less effect on the size of the movement. This is illustrated in Fig. 5, which shows the two possible positions for the pivot arrangement, namely positions A and B. When the pivot is at position A, only a small distance d1 exists between this pivot and the plane 60 (lying perpendicular to the page) of the afbre-mentioned interface area, whereas when the pivot is at position B, a much greater distance, d2, exists. Consequently, use of position B could give rise to much greater changes Ad in the position of the first stator blading 24 relative to the frame against which it is supposed to seal. Of course, in the case of the actual embodiment illustrated, the incorporation of the pivot at point B would not actually be feasible, since the transition duct 22 and guide vanes 24 could not be swung out of the engine. However, with some amendments to the layout, pivoting at B might be possible, though inadvisable due to the factors just discussed.

Ideally, the pivot arrangement should be close to the frame radially as well.

Referring again to Fig. 5, this means that it is preferable to minimise the distance d3 as much as possible for the same reason. This can be achieved, for example, by suitable design of the flange connecting the combustor to the engine frame. Thus, for instance, the flange could have an angled portion, which is bent over on the turbine side, allowing the pivot to be radially closer to the turbine blades.

Another way of achieving the same effect is shown in Fig. 6. In this arrangement two pivots are used to connect the combustor unit to the engine. One pivot 28 corresponds to pivot arrangement 28 of Fig. I and connects the combustor unit to a link 90, while the other pivot 29 connects the link 90 to the engine. This arrangement not only substantially meets the optimum conditions set out in Fig. 5, it also allows the combustor unit to swing further out from the engine opening, making it easier to access the components inside the engine. Depending on the engine design, it might be expedient to employ more than one link, and therefore more than two pivots.

The invention can be employed with combustors operating under a swirl burner principle, or indeed with any other type of combustor, provided it can be adequately supplied with fuel and air at the same time as being displaceably mounted to the engine such as to allow rapid and easy removal of critical components, as already explained.

Although the further seal has been described as a honeycomb seal, it may take any other suitable form. Furthermore, this seal could be mounted on the engine frame instead of on the end of the transition duct. However, securing the seal to the transition duct has the advantage that it is removed as part of the combustor unit, allowing easy inspection of the seal without having to interfere with other, internal parts of the engine.

These other types of seal could beneficially be used in conjunction with the "stop" principle (see stop elements 56 in Fig. 4C), especially where these types are insufficiently robust to accept the static seal mounting forces on their own.

Displacement mechanisms other than a pivot can be used. For example, a mechanism such as that shown in Figs 7A and 7B allows the modules to be temporarily attached to the engine frame before being clamped down or, alternatively, removed.

Fig. 7A shows a top view of a combustor 70, which is part of a combustor unit and is attached to a flange 72, which in turn is loosely placed against the engine frame 74 (see

II

Fig. 7B) by way of an intervening seal 76. The opening provided in the frame is, in this case, a rectangular opening 78, which allows the afore-mentioned "smile" to be easily removed. Attached to opposite sides of the flange 72 are two tabs 80, which extend from the flange sides first in a direction perpendicular to those sides, then angled in a direction parallel thereto. Adjacent the tabs are respective lugs 82 attached to the frame 74. To mount the flange and combustor unit, the flange is positioned so that the angled ends of the tabs 80 lie against the seal 76 with the tabs spaced apart from the end of the lugs 82. The flange is then slid laterally, so that the tabs become located under the angled part of the lags, which prevents (albeit temporarily) the flange and combustor unit from falling out of the opening 78. Finally, a clamp (as before, a band clamp) 84 is introduced over a raised portion 86 of the tabs of all the combustor units in the engine, and is tightened so as to secure the flanges of the units to the engine frame. A separate clamp is used for the two tabs. The reverse procedure is followed to remove the units from the engine.

A further alternative displacement mechanism involves the use of a spring, which attaches the flange to the engine casing. Thus the spring replaces the pivot and temporarily holds the combustion unit adjacent the engine opening prior to securing the unit to the engine or removing the unit from the engine.

Although it has been assumed that the clamps used will be a band clamp, other types of clamp may be used -for example a screw-down clamp (not shown), at least two for each combustor unit.

The above-described embodiments achieve a system for enabling rapid and convenient maintenance of mission-critical engine components, including those still within the engine when the combustor unit has been swung back or removed. Such maintenance may include scrubbing and vacuum-cleaning the inside of the engine all the way from the compressor exit to the turbine inlet. This is expedient, since in use deposits will tend to build up on the engine walls (this regardless of the air-inlet filtering which is employed) and there may also be some surface corrosion, which could break loose and pass through the turbine and cooling system, thereby causing damage. One engine component that may need only very occasional maintenance or replacement, for example, is the exit guide vane arrangement 18 (see Fig. 1). in practice, this is partially obscured by the compressor diffuser, which is conventionally used in gas turbine engines and converts the dynamic pressure of the compressed air leaving the compressor to static pressure, before this air is introduced into the combustor. Fig. 5 shows such a diffuser as item 62. If this diffuser is implemented as bolted-on arcuate segments, these segments can be readily removed, in the present invention, through the combustor opening, giving easier access to the exit guide vane arrangement 18.

A further advantage of the present invention, however, is that the combustor unit itself may be conveniently exchanged for a new or reconditioned unit. If such replacement units are obtainable aspre-assembled and pre-tested modules, then this can provide a further saving on service and commissioning time, during which the customer is deprived of the use of the engine. This benefit may prove commercially very attractive to a user, to whom down-time is a serious matter -for example a user in the oil industry. Moreover, if such units or modules are held locally as spares, then this can greatly save time in recommissioning an engine following an unexpected outage. It would also be possible to package such pre-tested modules with a consumable sealing material (e.g. a lightly waxed paper), which would be destroyed during initial operation without leaving any appreciable residue, which was detrimental to the module or other parts of the engine. This could reduce still further the risk of inadvertent contamination of the module on-site.

Claims

1. A gas turbine engine comprising a plurality of high-temperature modules, each module comprising a combustion unit for combusting a mix of a fuel and an oxidant, and a transition duct for conveying the combustion gases produced by the combustion unit to the rotor blading of the engine, wherein each module is secured to the engine by means of a displacement mechanism enabling rapid, easy access to the module and adjacent parts of the engine.

2. A gas turbine engine according to claim I, wherein the end of the transition duct is arcuate in shape, and an opening, which is provided in the engine for insertion or removal of at least parts of the module, is configured so as to allow insertion or removal of the transition duct.

3. A gas turbine engine according to claim 2, wherein each module further comprises stator blading disposed at the end of the transition duct for receiving the combustion gases from the transition duct, the module directing the gases to the rotor blading.

4. A gas turbine engine according to claim 3, wherein the stator blading is arcuate in shape, the opening being also configured so as to allow insertion or removal of the stator blading.

5. A gas turbine engine according to any one of claims 2 to 4, wherein the opening is rectangular or oval in shape.

6. A gas turbine engine according to any one of claims 2 to 5, wherein the displacement mechanism includes a pivot mechanism, which allows said at least parts of the module to be selectively inserted into or removed from the engine through the opening.

7. A gas turbine engine according to claim 6, wherein the pivot mechanism comprises a single pivot attaching the module to the engine.

8. A gas turbine engine according to claim 6, wherein the pivot mechanism comprises a first pivot attached to the engine, a second pivot attached to the module and at least one intermediate link between the first and second pivots.

9. A gas turbine engine according to any one of claims 2 to 5, wherein the displacement mechanism includes at least two lugs attached to the engine at opposite ends of the opening and a corresponding number of tabs attached to the module, whereby the module can be temporarily held adjacent the opening prior to selective securing of the module to the engine or removal of the module from the engine.

10. A gas turbine engine according to any one of claims 2 to 5, wherein the displacement mechanism comprises a spring, one end of which is attached to the engine and the other end of which is attached to the module, whereby the module can be temporarily held adjacent the opening prior to securing of the module to the engine or removal of the module from the engine.

11. A gas turbine engine according to any one of claims 6 to 8, wherein a clamp is provided at a part of each module, which, when the module is seated against the engine, lies on an opposite side of the opening to the pivot mechanism, thereby to retain the module in its seated position.

12. A gas turbine engine according to claim 11, wherein a further clamp is provided at, or close to, said pivot mechanism, thereby to force the module against the engine at the location of the further clamp.

13. A gas turbine engine according to claim 9 or claim 10, wherein at least one clamp is provided in contact with the module, thereby to force the module against the engine.

14. A gas turbine engine according to any one of claims 11 to 13, wherein the clamp, the at least one clamp and/or the further clamp is a band clamp, which co-operates with all of the modules simultaneously.

15. A gas turbine engine according to any one of claims 6 to 10, wherein each module comprises a flange, which is bolted to the engine casing.

16. A gas turbine engine according to any one of claims 3, 4, 7, 8 and II to 15, claim 5 as appendant to claim 3 or claim 4, and claims 6, 9 and 10 as appendant to any one of claims 3 to 5, wherein a seal is provided at an interface between a distal end of the stator blading and a part of the engine housing proximate to a first rotor blading of the engine.

17. A gas turbine engine according to claim 16, wherein said seal is secured to the distal end of the stator blading, being thereby removable as part of the module.

18. A gas turbine engine according to claim 17, wherein said seal is a honeycomb seal.

19. A gas turbine engine according to any one of claims 6 to 8, 17 and 18, wherein said pivot mechanism lies on a side of the module, which is located nearest the first rotor blading.

19. A gas turbine engine according to claim 18, wherein the honeycomb seal includes among its cells solid stop elements, which are configured such as to prevent the cells from being crushed in a sealing state of the honeycomb seal.

20. A gas turbine engine according to any one of claims 6 to 8, wherein the modules are connected to fuel manifolds, which are located on a pivot side of the modules.

21. A gas turbine engine according to claim 20, wherein the modules are connected to the fuel manifolds by way of flexible fuel lines, such that the fuel lines do not have to be disconnected when said at least parts of the modules are removed from the engine.

22. A gas turbine engine according to any one of claims 6 to 8, 20 and 21, wherein said pivot mechanism lies on a side of the module, which is located nearest the first rotor blading.

AMENDMENTS TO THE CLAIMS HAVE BEEN FILED AS FOLLOWERS

1. A gas turbine engine comprising a plurality of high-temperature modules, each module comprising a combustion unit for combusting a mix of a fuel and an oxidant, and a transition duct for conveying the combustion gases produced by the combustion unit to the rotor blading of the engine, wherein each module is secured to the engine by means of a displacement mechanism enabling rapid, easy access to the module and adjacent parts of the engine, wherein each module is mounted in an opening in a frame of the engine, and the operating action of the displacement mechanism is such as to remove the module from the opening to a position outside the engine frame, thereby allowing said easy access to the module and, via the opening in the frame, said easy access to the adjacent parts of the engine.

o 2. A gas turbine engine according to claim 1, wherein the end of the transition duct is (0 arcuate in shape.

3. A gas tuthine engine according to claim 2, wherein each module further comprises stator blading disposed at the end of the transition duct for receiving the combustion gases from the transition duct, the module directing the gases to the rotor blading.

4. A gas turbine engine according to claim 3, wherein the stator blading is arcuate in shape.

6. A gas turbine engine according to any one of claims 2 to 5, wherein the displacement mechanism includes a pivot mechanism. -by

9. A gas turbine engine according to any one of claims 6 to 8, wherein a clamp is provided at a part of each module, which, when the module is seated against the engine, lies on an opposite side of the opening to the pivot mechanism, thereby to retain the module in its seated position.

10. A gas turbine engine according to claim 9, wherein a further clamp is provided at, or close to, said pivot mechanism, thereby to force the module against the engine at the CD location of the further clamp.

11. A gas turbine engine according to claim 9 or claim 10. wherein the clamp and/or the further clamp is a band clamp, which co-operates with all of the modules simultaneously.

12. A gas turbine engine according to any one of claims 6 to 8, wherein each module comprises a flange, which is bolted to the engine casing.

13. A gas turbine engine according to claim 3 or claim 4, wherein a seal is provided at an interface between a distal end of the stator blading and a part of the engine housing proximate to a first rotor blading of the engine.

14. A gas turbine engine according to claim 13, wherein said seal is secured to the distal end of the stator blading. being thereby removable as part of the module.

15. A gas turbine engine according to claim 14, wherein said seal is a honeycomb seal.

16. A gas turbine engine according to claim 15, wherein the honeycomb seal includes among its cells solid stop elements, which are configured such as to prevent the cells from being crushed in a scaling state of the honeycomb seal.

17. A gas turbine engine according to any one of claims 6 to 8, wherein the modules are connected to fuel manifolds, which are located on a pivot side of the modules.

18. A gas turbine engine according to claim 17, wherein the modules are connected to the fuel manifolds by way of flexible fuel lines, such that the fuel lines do not have to be disconnected when the modules are removed from the engine.