GB2116639A - Turbine shroud segments and turbine shroud assembly - Google Patents

Abstract

A turbine shroud 15 for a gas turbine rotor blade stage 7 comprises shroud segments 25 made of a low strength metallic alloy of the "MCrAlY" type which is very highly oxidation resistant at elevated temperatures. The shroud segments 25 are supported by a structure in the form of a carrier ring 29 consisting of an alloy which retains high strength at high temperatures, and shroud cooling air chambers 53 are defined between the shroud segments 25, strengthening ribs (54 Fig. 2 not shown) of the shroud segments, and throttle plates 58 which also form part of the support structure for the shroud segments. Cooling air from chamber 49 is metered into shroud chambers 53 through holes 57 in throttle plates 58 so that the pressure in chambers 53 is only just enough to ensure exhaustion of the cooling air through holes 55 into the turbine passage just downstream of the rotor blades 7. This ensures not only low consumption of cooling air but also that the shroud segments are thrust outwards against their seatings on the throttle plates 58 by the higher turbine gas pressures upstream, the ribs (54) defining load paths which give the segments adequate distributed support to prevent overstressing or overstraining of the low strength alloy. <IMAGE>

Description

SPECIFICATION Turbine shroud segments and turbine shroud assembly The present invention relates to a metallic shroud segment and a shroud assembly for an axial flow gas turbine.

The present invention relates to a metallic shroud assembly for an axial flow gas turbine.

One of the factors affecting efficient operation of axial flow gas turbine aeroengines is the amount of cooling air which it is necessary to use in order to keep metallic turbine components operating at safe temperatures for the materials of which they are made. Because cooling air is extracted from the compressor (i.e. from an earlier part ofthethermo- dynamic cycle) and passed through turbine components, the work which it would have done in the turbine is largely lost, with deleterious effects on the power and specific fuel consumption of the aeroengine. Manufacturers are therefore anxious to reduce the amount of cooling airtaken by various turbine components without reducing the service life or safety of their engines.

One type of turbine component which has required a large amount of cooling air, is the metallic shroud ring surrounding the first or high pressure stage of turbine blades, the shroud ring being composed of a plurality of segments to allow for circumferential expansion and contraction due to temperature changes. This turbine shroud, like the turbine blades which it circumscribes, experiences high temperatures and pressures and thereforeagain like the turbine blades - has been a superalloy component requiring cooling with relatively large amounts of cooling air bled from the compressor.

It is therefore desirable to provide a metallic shroud assembly for a gas turbine rotor stage, which shroud assembly requires less cooling air relative to prior proposals.

Accordingly, one aspect of the invention contributes towards this goal by providing a turbine shroud segment composed of an alloy of the "MCrAIY" type, where M is a suitable major metallic constituent of the alloy and Y is the symbol for yttrium but also stands for any other suitable metallic element classed with the rare earths. Preferably, the major metallic constituent M of the alloy is iron and the element Y classed with the rare earths is yttrium or hafnium. It is also preferably that the alloy is oxide dispersion strengthened.

Such shroud segments can endure high temperatures with less oxidation than superalloys and hence require less cooling air in this respect, but although highly oxidation-resistant at high temperatures, MCrAIY-type alloys are of low strength relative to the superalloys normally used and therefore conventional methods of supporting, locating and cooling shroud segments are not applicable to MCrAIY-type alloys.

Accordingly, another aspect of the invention provides a turbine shroud assembly incorporating a plurality of the above-described shroud segments, this turbine shroud assembly including a shroud ring comprising said plurality of shroud segments, supporting structure circumferentially surround ing the shroud ring and to which the shroud segments are retained, the supporting structure consisting of a metallic alloywhich retains high strength at elevated temperatures, shroud chamber means defined between said shroud segments and said supporting structure, means for supplying cooling air from chamber means external to the assembly to pressurise said shroud chamber means, and means for exhausting cooling air from said shroud chamber means to a location downstream of the rotor stage;; said means for supplying cooling air to said shroud chamber means being adapted to meter said cooling air during operation of the turbine such that the total pressure forces acting outwardly on the shroud segments due to turbine gas pressure are substantially greater than the total pressure forces acting inwardly on the shroud segments due to cooling air pressure in said shroud chamber means, the shroud segments thereby experiencing an outward thrust, said shroud segments having means defining a plurality of load paths distributed over the radially outer sides of the shroud segments so as to transfer said outward thrust to said supporting structure without overstressing the shroud segments.

The outward thrust on the shroud segments can best be ensured by arranging that during operation of the turbine the pressure of the cooling air in the shroud chamber means is only just sufficient to ensure exhaustion of the cooling air to the location downstream of the rotor stage.

Other features of the invention will become apparent from the description of specific embodiments which follows and the appended claims.

An embodiment of the invention will now be described by way of example only with reference to the accompanying drawings, in which: Figure 1 is a "broken-away" sectional side elevation of part of a gas turbine incorporating the invention; Figure 2 is a view on section A-A in Figure 1.

The drawings are not to scale.

Referring in more detail to Figure 1, there is shown part of an axial flow gas turbine 1 as incorporated in a turbofan aeroengine. The turbine 1 has an annular turbine gas passage 3 in which are situated in flow series an annular array of nozzie guide vanes 5, a stage of turbine rotor blades 7, and an annular array of stator vanes 9, only the radially outer portions of these features being shown. Gases 11 from a combustion chamber exit 12 flow pastthe nozzle guide vanes 5, are guided thereby onto the turbine rotor blades 7, and from thence flow past the stator vanes 9 to the next stage of turbine blades (not shown).

The effective outer boundary of the illustrated portion of turbine gas passage 3 is formed by the outer shrouds 13 of guide vanes 5, a metallic turbine shroud ring 15, a flanged filler ring 17, and the outer shrouds 19 of stator vanes 9.

Guide vanes 5 and stator vanes 9 are fixed at their radially inner ends to static structure (not shown) is known manner. The forward ends of the outer platforms 13 and the inner platforms (not shown) of the guide vanes 5 locate against corresponding portions of the combustion chamber exit 12. Vanes 5 are additionally located at their radially outer platforms 13 against features on a frusto-conical drum member 21 as shown, and the forward parts of outer platforms 19 of vanes 9 are engaged with the rear edge of a support ring 23. Filler ring 17 is also held front and rear by support ring 23. Support ring 23 is itself connected to an outer casing (not shown) of turbine 1, as is the frusto-conical member 21.

The shroud ring 15 is provided to surround the radially outer tips of turbine blades 7 and form a seal againstthem in orderto prevent excessive leakage of the turbine gases over the blade tips between the high pressure and low pressure flanks of the blades.

It is composed of a number of shroud segments 25, which describe short arcs in the circumferential direction, this being illustrated in Figure 2. Sealing between adjacent shroud segments 25 against ingress of gas 11 through the gaps between adjacent segments is provided by means of so-called "stripseals" 27, which are well known to those skilled in the art, these being narrow metallic strips of relatively small thickness which are a clearance (sliding) fit in slots machined in circumferentially adjacent edges of the shroud segments. The shape of the slots and strip-seais 27 is shown in dashed lines in Figure land in cross-section in Figure 2.

The shroud segments 25 composing shroud ring 15 are retained to supporting structure which circumferentially surrounds the shroud ring. The supporting structure is an annular metallic carrier ring 29 and the shroud segments are retained to it by means of a "tongue-and-groove" or "hook" arrangement in which grooves 31 provided in the front and rear edges of the shroud segments engage respective rearwardly and forwardly projecting circular tongues 33 at the front and rear of the carrier ring, the shroud segments being a sliding fit between the tongues 33.

Carrier ring 29 is itself divided into a number of segments to allow for circumferantial expansion, these being of greater arc length than the shroud segments, e.g. each carrier ring segment holds three shroud segments. A split line between two carrier ring segments is shown at 34 in Figure 2. The carrier ring segments are held in position between support ring 23 at their rear and end-ring 35 of frusto-conical member 21 at their front. Support ring 23 is provided with a radially projecting annular flange 37, to which matching flanges 39 on the segments of carrier ring 29 are bolted.In order to support the front of the carrier ring 29 whilst allowing for relative movement due to thermal expansion and contraction, the front of the carrier ring segments are provided with forwardly projecting circular flanges 42 and the rear of the end ring 35 is provided with a circular siot 43, the flanges being received in the slot in a sliding fit as shown.

The outer sides of shroud segments 25 are provided with straight-sided depressions or cham bers 53 which are defined between strengthening ribs 54 extending fore-and-aft across circumferentially opposite ends of each of the segments to form a box-section as best seen in Figure 2. In the present embodiment the shroud segments 25 are of relatively short span in the circumferential direction, each requiring the support of only two ribs 54.

However, one or more extra ribs or pillars 54' (dashed lines) could be incorporated at equally spaced intervals across the span if necessary to provide extra support. We deem it desirable for the unsupported spans of the shroud segments to be small because of the limited high temperature strength of the alloy we utilise for the shroud segments.

Carrier ring 29 basically comprises ring sections front and rear for connection to neighbouring structure as already mentioned, and a cylindrical section connecting the front and rear ring sections, the cylindrical section being provided with large circumferentially spaced apertures 56. Sandwiched between the carrier ring 29 and the shroud segments 25 are part-cylindrical throttle plates 58 which in this case are substantially coextensive axially and circumferentially with the shroud segments, though they could Id be circumferentially coextensive with the carrier ring segments.Carrier ring 29 and shroud segments 25 are designed to receive the throttle plates between them, and the throttle plates are held against sliding movement relative to the carrier ring 29 by location pins (not shown) which protrude from the carrier ring into corresponding holes in the throttle plates. However, the throttle plates are not substantially restrained to the carrier ring 29 in the radially inward direction. It should be noted that throttle plates 58 make contact with carrier ring 29 only over narrow strips near their front and rear edges, but that they make contact with shroud segments 25 not only over the narrow strips near their front and rear edges, but also over the outer surfaces of ribs 54. These ribs 54 therefore provide a seal against the throttle plates 58.

Cooling air for stator vanes 9, carrier ring 29 and shroud segments 25 is supplied as indicated by the arrows 45 from annular chamber 47 surrounding frusto-conical member 21. Chamber 47 is fed by an air bleed from the compressor (not shown) of the turbofan. Chamber 47 communicates freely with chamber 49 surrounding carrier ring 29, and chamber 49 supplies chamber 51 surrounding the outer plafforms 19 of vanes 9. Stator vanes 9 are hollow and require cooling with air from chamber 51 as shown. In order to cool shroud segments 25, cooling air from chamber 49 flows through apertures 56 in the carrier ring 29 (causing slight cooling of the same) and enters shroud chambers 53 on the outer sides of the shroud segments after being metered through small holes 57 in the throttle plates 58. The cooling air is exhausted from the chambers 53 into the turbine passage 3 just downstream of the turbine blades 7 by means of angled drillings 55 through the rear edges of the segments 25.

Particular reference will now be made to features in the design which facilitate economic use of the cooling air.

Although actively cooled as mentioned above, the alloy of which the shroud segments 25 are made has a high resistance to oxisation at high temperatures.

In designs for known types of metallic shroud segments made from superalloys, the temperatures of the shroud segments are kept within acceptable upper iimits by supplying large mass flows of cooling air to the segments for subsequent exhaustion to the turbine passage. However, our use of the more highly oxidation resistant alloys allows higher metal temperatures to be tolerated in the shroud segments without increased danger of failure due to stress concentrations caused by oxidation of the metal, hence the shroud segments require less cooling air and the efficiency of the engine can be increased.

In the present instance it is desired to run the shroud segments at temperature in excess of 1100 C on their inner surfaces. We have found that an yttria dispersion strengthened FeCrAIY alloy or a hafnia dispersion strengthened FeCrAlHf alloy highly oxidation resistant for this purpose. Hitherto FeCrAIYtype alloys have been known for use as elements in electric furnaces, and as highly oxidation-resistant coatings for machine components made of other less oxidation-resistant alloys, such as nickel-base superalloy gas turbine blades, etc. Other highly oxidation resistant alloys of this general type are known, such as CoCrAIY and NiCrAIY alloys, these being generically referred to as "MCrAIY" alloys, where M is a suitable major metallic consistituent of the alloy as known to those skilled in the art.We prefer to use the dispersion strengthened FeCrAIY or FeCrAIHf alloys because they have a higher softening temperature than other MCrAIYtypes, but other MCrAIY types could be used if the correct balance between the heating effect of the turbine gases on the shroud segments and economical use of cooling air is achieved in each case.

It is possible that suitable metallic oxides other than yttria or hafnia, classed with the rare earth oxides, could be used to strengthen the alloy chosen for the shroud segments. Note that it is necessary to produce such alloys for machine components from powder materials by means of a mechanical alloying process as known to those skilled in the art.

As an example, a basic FeCrAIY alloy useful for putting the invention into effect has the composition.

Carbon < .03% Chromium 15-20% Aluminium 4-5% Yttrium 0.05-0.4% Iron the rest.

A problem associated with the use of MCrAIY-type alloys for structural members such as the shroud segments 25 is their very low ultimate tensile strength (UTS). Whereas a typical superalloy may have a UTS of about 48 x 107 Pa, the FeCrAIY alloy used here may have a UTS of only about 0.8 x 107 Pa.

The invention is thus intimately concerned with ensuring that the forces experienced by the MCrAIY shroud segments during operation of the turbine do not exceed the strength of the alloy, these forces being those due to the pressures of the cooling air and the turbine gases. This statement will be amplified by considering the balance of forces across the shroud segments 25 in Figures 1 and 2.

in the illustrated arrangement, the only important radially inward pressure forces on each shroud segment are: i) the force due to the pressure in chamber 49 acting on the solid area of throttle plate 58 exposed to that pressure (N.B. the throttle plate is free to thrust radially inward against the shroud segments); and ii) the force due to the pressure of the cooling air in shroud chambers 53 acting on the radially inner surfaces of the chambers.

The only important radially outward pressure force on each shroud segment is the force due to the pressure which the turbine gases exert on the radially inner face of the shroud segment. This pressure varies between the front and rear edges of the shroud segments, the pressure just upstream of the row of blades 7 being much greater (by a factor of 1.5-2.0) than the pressure just downstream of the row. Pressures at intermediate positions on the inner faces of the shroud segments are (when averaged out between high pressure and low pressure flanks of the blades) intermediate in value.

It is an important result of the present invention that even though the sum of the above radially inward forces i) and ii) may actually exceed the radially outward force by a large amount, the radially inwardly unsupported span of each shroud segment 25 (i.e. the part extending between the front and rear tongue-and-groove engagements with the carrier ring 29) experiences only a net outwardly directed thrust of pressure force which causes ribs 54 to bear outwards and react against throttle plates 58, thereby defining load paths which give the segments adequate distrubuted support against the bending effects of the outwardly directed pressure force so as to prevent overstressing or overstraining of the segments. Moreover, when ribs 54 bear outwards against the throttle plates, shroud chambers 53 are sealed against entry of turbine gases should any get past the strip seals 27.

Remembering that the shroud segments comprise a low strength (and hence low rigidity) material, this desirable result is brought about in the present embodiment by making the throttle plates 58 from a high-strength, highly rigid material which retains its strength at high temperatures, such as a nickelbased superalloy. Thus, the throttle plates are substantially rigid relative to the shroud segments under the pressures involved and they span the distance between the tongue and groove features such that the inward pressures forces on the plates are transmitted straight through the front and rear outer edge portions of the shroud segments as compressive loads for reaction against the tongues 33 of the carrier ring 29, which is also made of a superalloy.

By this means, the shroud segments do not experience any bending effect from inwardly directed pressure forces due to the pressure in chamber 49, but only the bending effects of the inward pressure force due to the cooling air in chambers 53 and the outward pressure force due to the turbine gases 11.

Consequently, in order to achieve the desired result of a net radially outward pressure force acting on each shroud segment, it is arranged that the pressure of the cooling air in the chambers 53 on the outer sides of the shroud segments 25 is only just sufficient to ensure adequate exhaustion of the cooling air to the turbine passage 3through drillings 55, i.e. the pressure in chambers 53 is only slightly higher than the pressure of the turbine gases at the rear edges of the segments just downstream of the turbine blades 7.Because the pressure of the turbine gases on the more forward regions of the shroud segments is greater than it is near their rear edges, the segments experience an outwardly acting pressure force from the turbine passage which is greater than the inwardly acting pressure force due to the pressure of the cooling air, thereby causing the segments to be thrust outwardly against their seatings on the throttle plates as required.

The supply pressure of the cooling air in the chamber 49 can of course be the same as for previous designs of shroud segments because the cooling air 45 is required for other tasks such as cooling statorvanes 9. The required metering of the cooling air supply to the shroud segments, i.e. the required drop in pressure between chamber 49 and chambers 53, is achieved by careful sizing and spacing of holes 57 in throttle plates 58.

It will be noted that carrier ring 29 and throttle plates 58 are shielded from the direct effects of the hot gases 11 by the shroud segments 25, and they also experience some cooling due to the flow of cooling air into chambers 53 of the shroud segments. However, the conductive transfer of heat into these components from the shroud segments 25 can also be minimised by providing, at the iocations where the shroud segments make contact with the carrier ring 29 and the throttle plates 58, a thermal barrier coating on the shroud segments and/or on the carrier ring and the throttle plates. Known thermal barrier coatings include, for example, boron nitride, yttria stabilized zirconia, or the so-called "magnesium zirconate" materials available from such manufacturers as Metco.

The forward inner edge of the carrier ring 29 is conventionally protected from the effects of turbine gases 11 entering the gap between the rearward edges of the vane platforms 13 and the forward edges of the shroud segments 25, by means of high pressure air 59 which is fed to the gap via drillings 63 and clearance 64 from annular chamber 61 between platforms 13 and frusto-conical member 21. The air 59 is supplied to the gap at a pressure in excess of the pressure of turbine gases 11 just upstream of the turbine blades 7, the chamber 61 being pressurised buy a bleed from the compressorto a pressure considerably in excess of the pressure in chamber 47.

Similarly, the rear inner edge of the carrier ring 29 is protected from turbine gases 11 by means of air 65 which is fed to the gap between the rear edges of the shroud segments and the forward edge of the filler ring 17 from chamber 49 via drillings 67 and clearance 68. The air 65 can be supplied at a lower pressure than air 59 because of the lower pressure of the turbine gases 11 downstream of the turbine blades 7.

In Figures 1 and 2, the ribs 54 on the radially outer sides of segments 25 make contact with the radially inner surfaces of the throttle plates 58 in order to define load paths for transferring the radially outward pressure forces on the segments to the carrier ring 29. In an alternative arrangement (not shown), the radially outer sides of the shroud segments make contact with support structure through load paths defined by areas of contact between ribs provided on the shroud segments as before, and further ribs provided on the support structure, the further ribs being oriented transversely of ribs on the shroud segments. The ribs on the support structure may be on throttle plates provided as separate components sandwiched between the carrier ring and the shroud segments as in Figures 1 and 2.

Alternatively, throttle plates as separate components may be absent, the ribs being provided on a radially inner surface ofthe carrier ring. In either case, means are provided for throttling the supply of cooling air to the chambers between the ribs on the shroud segments on the same principie as explained in relation to Figures 1 and 2. Note that if the cooling air throttling function is performed by holes in an integral portion of the carrier ring, rather than by separate throttle plates, the shroud segments do not have to cope with the radially inward pressure forces transmitted by such throttle plates.

The provision of ribs on the support structure as well as on the shroud segments produces smaller areas of contact between the support structure and the shroud segments, thereby reducing conductive heat transfer from the shroud segments to the support structure. Heat transfer may be further reduced by the use of thermal barrier coatings as previously described. In orderto provide for greater cooling of the support structure and the shroud segments, the holes which feed cooling air to the chambers between the ribs on the shroud segments -may extend as drillings through the ribs on the support structure and through the ribs on the shroud segments, these drillings acting to cool both sets of ribs before discharging to the chambers between the ribs.

Note that in the case ofthe embodiment described in relation to Figures 1 and 2 above, and in the case of the alternative embodiment described above, cooling of the shroud segments can be enhanced without necessarily using more cooling air by ensuring that cooling air supplied to the chambers between the ribs on the shroud segments issues from the cooling air holes or drillings in such a way as to form jets of cooling air which impinge on the radially inner sides of the chambers to pierce the boundary layer of hot air and hence cool the shroud segments more effectively.

Claims (17)

1. A turbine shroud segment for a gas turbine rotor blade stage, the shroud segment being composed of an alloy of the "MCrAIY" type, where M is a suitable major metallic constituent of the alloy and Y is a metallic element classed with the rare earths.

2. A turbine shroud segment according to claim 1, the major matallic constituent M of the alloy being iron.

3. Aturbine shroud segment according to claim 1 or claim 2, the element Y classed with the rare earths being yttrium or hafnium.

4. Aturbine shroud segment according to any one of claims 1 to 3, the alloy being oxide dispersion strengthened.

5. Aturbine shroud assembly incorporating a plurality of shroud segments according to any one of claims 1 to 4, the shroud assembly including a shroud ring comprising said plurality of shroud segments, supporting structure circumferentially surrounding the shroud ring and to which the shroud segments are retained, the supporting structure consisting of a metallic alloy which retains high strength at elevated temperatures, shroud chamber means defined between said shroud segments and said supporting structure, means for supplying cooling air to pressurise said shroud chamber means, and means for exhausting cooling air from said shroud chamber means to a location downstream of the rotor stage;; said means for supplying cooling airto said shroud chamber means being adapted for metering said cooling air during operation of the turbine such that the total pressure forces acting outwardly on the shroud segments due to turbine gas pressure are substantially greater than the total pressure forces acting inwardly on the shroud segments due to cooling air pressure in said shroud chamber means, the shroud segments thereby experiencing an outward thrust, said shroud segments having means defining a plurality of load paths distributed over the radially outer sides of the shroud segments so as to transfer said outward thrust to said supporting structure without overstressing the shroud segments.

6. Aturbine shroud assembly according to claim 5 in which the means for supplying cooling air to pressurise the shroud chamber means is adapted to meter said cooling air during operation of the turbine such that the pressure in said shroud chamber means is only just sufficient to ensure exhaustion of said cooling air to the location downstream of the rotor stage.

7. A turbine shroud assembly according to claim 5 or claim 6 in which the shroud chamber means comprise depressed portions of the radially outer sides of the shroud segments, said depressed portions being defined between features on said outer sides of the shroud segments, which features define the load paths between the shroud segments and the supporting structure.

8. A turbine shroud assembly according to claim 7 in which the features on the outer sides of the shroud segments comprise rib portions extending thereacross to strengthen said segments.

9. A turbine shroud assembly according to any one of claims 5 to 8 in which the supporting structure includes a surface for reacting the outward thrust from the shroud segments, and for sealing therewith, which surface is substantially cylindrical about the rotational axis of the turbine.

10. A turbine shroud assembly according to any one of claims 5 to 9 in which tongue features on the supporting structure engage groove features in mutually opposite edge portions of the shroud segments, whereby the shroud segments are retained to the supporting structure.

11. A turbine shroud assembly according to claim 5 in which each shroud segment has opposed edge portions defining groove features therein, the supporting structure including plate members defining a surface for reacting the outward thrust from the shroud segments and for sealing therewith, said plate members having holes therein for the metering of the cooling air, a ring portion outboard of the plate members, and hook means extending from the ring portion to define tongue features thereon, which tongue features engage said groove features, whereby the shroud segments are retained to the supporting structure;; wherein the plate members are sandwiched between the shroud segments and the ring portion of the supporting structure, the plate members spanning the distance between said opposed edge portions of the shroud segments such that at least some of the inward pressure forces on the plate members due to the metering of cooling air are transmitted through said opposed edge portions of the shroud segments for reaction against said tongue features.

12. A turbine shroud assembly according to any one of claims 5 to 10 in which the means for supplying cooling air to the shroud chamber means comprises a plurality of holes extending through the supporting structure, the holes being so sized and spaced-apartfrom each other as to meterthe cooling air into the shroud chamber means in the specified way.

13. Aturbine shroud assembly according to any one of claims 5 to 12 in which locations where the shroud segments make contact with the supporting structure are provided with thermal barrier coatings on at least the shroud segments.

14. Aturbine shroud assembly according to claim 13 in which the thermal barrier coating comprises boron nitride or "magnesium zirconate" or yttria stabilised zirconia.

15. Aturbine shroud segment substantially as described in this specification with reference to and as illustrated by the accompanying drawings.

16. A turbine shroud assembly substantially as described in this specification.

17. A turbine shroud assembly substantially as described in this specification with reference to and as illustrated by the accompanying drawings.