GB2401655A

GB2401655A - A rotor blade arrangement

Info

Publication number: GB2401655A
Application number: GB0311159A
Authority: GB
Inventors: Andy Che-Yeung Au; Tony David Pheaton
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2003-05-15
Filing date: 2003-05-15
Publication date: 2004-11-17
Also published as: GB0311159D0

Abstract

A rotor blade arrangement 31 is provided in which a number of aerofoils are secured through a common platform 32 and a more limited number of fixing roots 34. Thus rotor blade arrangements 31 can replace a number of individual blades secured to a rotor disc to form a rotor blade ring assembly reducing assembly time, machining and pro rata component weight. The arrangements 31, (61, Fig 6) may utilise existing fixture aperture structures in the rotor disc in order that a present arrangement 31, (61, Fig 6) can be utilised in an existing gas turbine engine assembly by replacing individual blades. Alternatively, specific fixing roots (44, Fig 4), (54, Fig 5) may be provided within reciprocal fixing apertures in a rotor disc. The rotor blade arrangement 31 is primarily of use in a compressor of a gas turbine engine.

Description

A Rotor Blade Arrangement The present invention relates to rotor blades

and more particularly to compressor blades used in compressors of gas turbine engines.

Referring to Fig. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust nozzle 19.

The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow 0 directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high pressure compressor 14 is directed into the combustion equipment 15 i- 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 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbine 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13, and the fan 12 by suitable interconnecting shafts.

The present invention particularly relates to compressor blades used in the compressors 13, 14. As can be seen, a number of compressor blades are arranged in a cascade in order to create the desired airflow compression for gas turbine engine operation. Fig. 2 is a front perspective view of a current compressor blade assembly 20 comprising a compressor disc 21 and compressor blades 22.

As can be seen, individual compressor blades 22 comprise an aerofoil 23, a platform 24 and a root 25 for securing the compressor blade 22 to the compressor disc 21. In Fig. 2 this means of securing the compressor blade 22 to the compressor disc 21 is through a firtree root 25. It will be understood that manufacture of each individual compressor blade 22 is expensive along with provision of respective firtree slots 26 in the compressor disc 21 in order to accommodate and secure those compressor blades 22 through the root 25. In addition, it will be appreciated that there is significant complexity with regard to replan ng each individual compressor blade 22 during maintenance or servicing.

According to the present invention there is provided a rotor blade arrangement comprising a plurality of aerofoils secured. to a common platform and that common platform including fixing means for collectively fixing that common platform and those aerofoils upon a rotor.

Typically, the fixing means comprises one pepen.iicular root fixing extending into the rotor.

Altern lively, the fixing means comprises two or more perpendicular root fixings penetrating into the rotor.

Typi a ly, the penetrating root fixings are of a firtree type o,- of a dovetail type.

Generally, the fixing means is positioned upon the cony n platform to mate with existing fixing features of the -or or specified for fixing individual rotor blades. i

Advantageously, the fixing means are directly secured to the common platform.

The rotor may be a disc or a drum.

Also in accordance with the present invention there is provided a rotor blade assembly comprising a plurality of rotor blade arrangements as described above with ends of the adjacent platforms abutting one another to form segments of a ring about the rotor.

In addition the present invention includes a gas turbine engine including a rotor blade arrangement and/or an assembly as described above.

The rotor blade arrangement may be a compressor blade arrangement or a turbine blade arrangement.

Embodiments of the present invention will now be described by way of example only with reference to the accompanying drawings in which: Fig. 3 is a schematic front view of a first embodiment of the present invention; Fig. 9 is a front schematic illustration of a second () embodiment of the present invention; Fig. 5 is a schematic front view of a third embodiment Of the present invention; and Fig. 6 is a schematic front view of a fourth embodiment of the present invention.

Rotor blades are generally aerofoils arranged upon a rotor disc, or a rotor drum, which rotates in order to generate airflow. The present invention particularly Relates to compressor assemblies used in a gas turbine engine in order to generate airflows within that engine.

O Typically, a number of compressor blade ring assemblies are arranged in cascade in order to generate airflow as depicted in Fig. 1. Traditionally, as depicted in Fig. 2 each compressor blade is secured through its own platform arid root mounting into a compressor disc. In such circumstances, each compressor blade and compressor disc must be machined for appropriate assembly of the compressor blades into the compressor disc and it will be understood that the gaps between adjacent platforms are susceptible to airflow ingress creating leakage problems and turbulence.

The present invention provides rotor blade arrangements in which a number of aerofoils are secured upon a common platform in order to reduce the number of abutments between adjacent platform edges and therefore leakage along with turbulence. It will also be understood that a reduced number of assembly steps are needed when several aerofoils are secured in one procedure rather than when each blade aerofoil is individually secured. Although more machining will be required upon the rotor blade arrangement than each individual previous rotor blade, it will be understood there is less collective machining per rotor blade by utilising rotor blade arrangements in accordance with the present invention. The present invention may also utilise securing the rotor blade arrangement within previous root fixing aperture structures for individual rotor blades such that upon maintenance and refurbishment those individual rotor blades can be replaced by cJ rotor blade arrangement in accordance with the present invent iC''1.

Fic. 3 illustrates a front view of a first embodiment of the present invention. Thus, a compressor blade arrangement 31 comprises a platform 32 upon which aerofoils 33 are secured with appropriate spacing and positioning.

The platform 32 has axial fixing roots 34 directly secured on an opposite side to the aerofoils 33. As shown in Fig. 3 tile f xing roots 34 are directly below respectively two of t-.ke aerofoils 33 and so are positioned where those aero1Gi s 33 if individual blades would have their fixing l root in a compressor disc of a compressor blade ring assembly.

The compressor blade arrangement 31 is typically cast in an appropriate material and then machined in order to achieve the desired final shaping for the aerofoil 33, platform 32 and in particular the fixing roots 34. As can be seen, the platform 32 is substantially arcuate and therefore in use will be associated with other compressor blade arrangements 31 in order to create a ring which is an assembly as part of the compressor (Fig. 1). End portions of each platform 32 abut with each other in order to form the ring. However, as described previously, these abutments between the end portions 35 create potential leakage problems and so by providing less joint abutments diminishes these problems.

Upon assembly, the compressor blade arrangements 31 will be presented to their appropriate reciprocal fixing structures within the compressor disc of the compressor assembly. As described previously, these reciprocal fixing -0 structures may be those originally configured for individual assembly of compressor blades in accordance with previous arrangements. In such circumstances, the present invention in the first embodiment depicted in Fig. 3 allows installation of four aerofoils 33 in a single operation.

As indicated previously, the fixing roots 34 in the first embodiment depicted in Fig. 3 are substantially below the outer aerofoils 33 and are consistent with the position in a compressor disc for the fixing root of those blades if individually formed and presented. These roots 34 are 0 substantially perpendicular to the arcuate platform 32 in order to achieve best presentation of the platform 32 and aerofoils 33. In such circumstances, the inner aerofoils 33 in the first embodiment have no fixing root 34 directly below them but are still resiliently presented as part of the compressor blade arrangement 31 through fixing roots 34. Provision of less fixing roots reduces weight as well as machining and assembly costs.

Fig 4 is a schematic front schematic illustration of a compressor blade arrangement 41 which comprises a platform 42, aerofoils 43 and root fixing 44. The platform 42 accommodates three aerofoils 43 with the single fixing root 44. The platform 42 is arcuate and as described previously will form a compressor ring with other compressor blade arrangements 41 in which the aerofoils 43 act to generate airflow and normally air compression. In order to resiliently anchor and fix the compressor blade arrangement 41 the fixing root 44 is generally more substantial than that described previously with regard to fixing roots 34 in Fig. 3. Nevertheless, by providing a single fixing root 44 it will be appreciated that assembly of the compressor blade arrangement 41 into a compressor blade assembly is made easier by the more limited number of fixing roots 44 to compressor disc associations required. Similarly, machining the arrangement 41 is easier than machining individual blade arrangements. As will be noted, the fixing root 44 is generally centrally located upon the platform 42 for balance in terms of fixing distribution and presentation of the aerofoils 43.

Fig. 5 illustrates a schematic front view of a third embodiment of the present invention. Thus, a compressor blade arrangement 51 includes a platform 52, aerofoils 53 and a circumferential fixing root 54. The platform 52 is again actuate and in association with other arrangements 51 will form a compressor ring assembly presenting the aerofoils 53 within a gas turbine engine. In a similar fashion to the embodiment depicted in Fig. 4 the circumferential fixing root 54 is centrally located with regard to the aerofoils 53 to provide a balanced location and fixing for the arrangement 51. Typically, a fixing band may be secured through an aperture 55 in the fixing route 54 to act in the direction of arrowheads 56 to draw together all the arrangements 51 in a compressor blade assembly.

Fig. 6 is a front view of a fourth embodiment of the present invention. Thus, the compressor blade arrangement 61 comprises a platform 62 with aerofoils 63 and fixing roots 64. The fixing roots 64 are of a circumferential type whereby through apertures 65 a fixing band or other means is arranged as shown by broken line 66 to draw the arrangement 61 into association with other arrangements 61 in order to define a compressor blade ring assembly for use within a gas turbine engine.

In a similar fashion to the first embodiment described with regard to Fig. 3 the fixing roots 64 are generally located below a particular aerofoil 63 and so will normally be shaped and configured to associate with a normal ircumferential fixing root aperture structure for an 0 individual blade in those positions. Thus, the arrangement 61 can replace three individual blades with a single arrangement 61 reducing assembly time. There is also a reduction in the weight of a fixing root below the central aerofoil 63 which is no longer required and a reduction in the machining required pro rata for the three aerofoils in comparison with three individual blades in a previous arrangement.

At the core of the present invention is the facility of providing assembly of a number of rotor blades more -onveniently in a rotor blade assembly. The present blade arrangements comprise segments of that rotor blade assembly secured to a rotor disc or a rotor drum, in order to -omprise a rotor blade ring bank in a gas turbine engine.

3v using rotor blade arrangements in accordance with the present invention the number of components which must be secured to each rotor disc or rotor drum is reduced which improves installation and assembly time. As indicated previously each blade/aerofoil was generally secured by its own fixing root and slot within the rotor disc and this significantly increased assembly time.

In accordance with the invention, the rotor blade arrangements can be accommodated within existing rotor discs or rotor drums utilising existing fixing apertures of those rotor discs or rotor drums through appropriate fixing roots 34, 64 or by provision of specific fixing in those rotor discs or rotor drums for use with regard to generally oversized fixing roots 44, 54 in comparison with previous fixing roots. In either event, the present rotor blade arrangement reduces assembly time and therefore cost of manufacture of a gas turbine engine incorporating a rotor stage comprising a number of rotor blade assembly rings or rotor blade banks. It will also be understood that a reduced number of fixing root to fixing apertures within the rotor disc reduces the number of machining operations in order to provide the accurate reciprocal portions to effectively mount rotor blades.

Previously, each rotor blade as indicated had its own platform which must be appropriately sized to ensure a minimum gap between adjacent rotor blades to prevent leakage and also must have an appropriate fixing root accurately machined for appropriate location in the rotor disc with little likelihood of detachment or wobble.

Reducing these machining processes will also reduce assembly and manufacturing costs.

Typically, in addition to the common platforms to which the fixing roots are secured an outer platform may also be secured to the aerofoil at the tips of those aerofoils. Such an outer platform secured to the tips of the aerofoils will resist flexing and deformation of the aerofoils in use as a result of centrifugal forces upon rotation about the rotor disc. In any event, the end 35 abutments between adjacent arrangements in accordance with the present invention will act to prevent such flexing in association with the material thickness of the common platform which may be increased as a result of the reduced weight by eliminating individual fixing roots for each rotor blade. Such thickening of the platform may also avoid problems with flexing of the platform due to reduced anchoring provided by a lack of individual fixing roots for each rotor blade.

As indicated previously, fixing roots in accordance with the present invention will generally be perpendicular to the arcuate platform in order to act appropriately to fix the arrangements in accordance with the present invention. Such fixing roots will also be generally balanced within a rotor blade arrangement to prevent wobble of the arrangement about the respective route fixings.

The rotor blade arrangements are primarily applicable to compressor blade arrangements but are applicable to turbine blade arrangements and are equally applicable to rotor discs and rotor drums. The fixing roots may be of the dovetail type or firetree type and may be of the circumferential type or axial type.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims

1. A rotor blade arrangement comprising a plurality of aerofoils secured to a common platform and that common platform including fixing means for collectively fixing that common platform and those aerofoils upon a rotor.

2. An arrangement as claimed in claim 1, wherein the fixing means comprises one perpendicular root fixing extending into the rotor.

3. An arrangement as claimed in claim 1, wherein the fixing means comprises two or more perpendicular root fixings penetrating into the rotor.

4. An arrangement as claimed in claim 2 or claim 3, wherein the or each penetrating root fixing is of a firtree type.

5. An arrangement as claimed in any preceding claim wherein the fixing means is positioned upon the common platform to mate with existing fixing features of the rotor specified for fixing individual rotor blades.

6. An arrangement as claimed in any preceding claim wherein the fixing means are directly secured to the common platform.

7. An arrangement as claimed in any preceding claim wherein the rotor is a disc or a drum incorporating the common platform.

8. A rotor blade arrangement substantially as hereinbeore described with reference to Fig. 3.

9. A rotor blade arrangement substantially as here nbeore described with reference to Fig. 4.

10. A rotor blade arrangement substantially as hereinbe ore described with reference to Fig. 5.

11. A rotor blade arrangement substantially as here nbeore described with reference to Fig. 6.

12. A turbine rotor blade assembly comprising a rotor blade arrangement as claimed in any preceding claim.

13. A compressor blade assembly as claimed in any preceding claim.

14. A rotor blade assembly comprising a plurality of rotor blade arrangements as claimed in any of claims 1 to 10 wherein ends of adjacent platforms abut one another to form segments of a ring about the rotor.

15. A gas turbine engine including a rotor blade arrangement as claimed in any of claims 1 to 10.

16. A gas turbine engine including a rotor blade assembly as claimed in claim 13.

17. Any novel subject matter or combination including novel subject matter disclosed herein, whether or not within the scope of or relating to the same invention as any of the preceding claims.