GB2412699A

GB2412699A - Heat shield for rotor blade hub

Info

Publication number: GB2412699A
Application number: GB0407072A
Authority: GB
Inventors: Alec George Dodd
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2004-03-30
Filing date: 2004-03-30
Publication date: 2005-10-05
Also published as: GB0407072D0

Abstract

A rotor blade arrangement 100, has a damper member 106 below platform sections 103, 104 both to prevent vibration chatter between adjacent rotor blades 101, 102, and also to provide a heat radiation barrier. This heat radiation barrier within the member 106 reduces heat radiation flux within a cavity 113 so a rotor disc 108 can be kept at an acceptable operational temperature. The heat radiation feature takes the form of a shield 201 secured upon base or wing sections 204, 205 of the damper member 106. As the space between adjacent blade roots 202, 203 is irregular, the heat shield feature 201 is also irregular to ensure damping function is maintained whilst achieving a heat flux radiation barrier.

Description

24 1 2699 Blade Mounting Arrangement The present invention relates to

blade mounting arrangements and more particularly to blade mounting arrangements used with regard to turbine blades used in a turbine engine.

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, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.

The gas turbine engine 10 operates 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 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 combustor 15 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 turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.

It will be understood from above that the turbine blade stages 16, 17, 18 are subject to high gas flow i temperatures. Furthermore, engine 10 efficiency is improved with higher gas temperatures such that it is desirable to operate the engine 10 at as high gas flow temperatures as possible. Clearly, in such a situation a major limitation is the melting point of the materials from which on particularly turbine components of the engine 10 are made. In such circumstances normally coolant air is taken from the compressor side of the engine 10 and utilised in the turbine side of the engine in order to provide cooling to appropriate parts.

A particular problem relates to the mounting of the turbine blades 16, 17, 18 upon rotor discs secured about the shafts of the engine 10. There are a number of approaches to securing the blades upon the rotor disc, including so called fir tree engagement between reciprocally shaped blade root segments and cavities within the rotor disc. Typically the blades are secured in close juxtaposed position to extend radially from the rotor disc with a damper member between adjacent blades utilised in order to facilitate secure location, a gas seal and vibration chatter limitation between blade platform segments extending into juxtaposed positions circumferentially in order to create the turbine stage in the engine 10. Nevertheless, it will be appreciated that the turbine blades as indicated are subject to high combustion gas temperatures such that there can be significant heating of the turbine disc in which the blades are secured.

In accordance with the present invention there is provided a turbine blade arrangement comprising a rotor disc upon which rotor blades are secured, a damper member between root sections of respective adjacent rotor blades and the damper member having a heat shield feature to inhibit heat transfer towards the rotor disc through blade platform sections extending between the adjacent rotor blades.

Preferably, the heat shield feature is integral with the damper member.

Typically, the heat shield feature comprises a preferential heat conduction element within the damper.

Alternatively, the heat shield feature comprises appropriate heat reflective coatings upon the damper member. Additionally, the heat shield feature may comprise appropriate shaping of the damper member to facilitate heat dissipation.

Possibly, the damper and heat shield feature are particularly configured for dynamic balance within a turbine blade arrangement for appropriate performance when that turbine blade arrangement is rotated. Possibly, such configuration of the damper and heat shield feature comprises a hollow section within the damper member appropriately shaped for dynamic balance within the turbine blade arrangement.

Generally, the damper member comprises a base section arranged to extend between root sections of the adjacent rotor blades and a radial rib upstanding from the lateral base section to engage the blade platforms. Typically, the rib engages at a juxtaposed joint between adjacent platform sections. Normally, the radial rib comprises a plurality of spikes or ridge extending along the lateral base member.

Also in accordance with the present invention there is provided a turbine blade assembly comprising a plurality of turbine blade arrangements as described above. Generally, that assembly incorporates coolant flow means and that coolant flow means is arranged to be directed in use towards the damper member in order to facilitate the heat shield feature.

Also in accordance with the present invention there is provided a turbine engine incorporating a turbine blade arrangement and/or a turbine blade assembly as described above.

Additionally, in accordance with the present invention there is provided a damper member for a turbine blade arrangement, the damper member comprising a heat shield feature to inhibit in use heat transfer across a turbine blade arrangement.

Typically, the heat shield feature comprises a low emissivity coating or element or shaping of the damper member. Furthermore, the damper member may have a hollow passage or cavity to provide in use dynamic balance for the damper member.

Embodiments of the present invention will now be described by way of example and reference to the accompanying drawings in which; Fig. 2 is a schematic cross-section of a rotor blade arrangement in accordance with the present invention illustrating two rotor blade roots in adjacent location; Fig. 3 is a schematic illustration of a damper located between adjacent rotor blades in accordance with the present invention; and Fig. 4 is a schematic cross section of a damper in accordance with the present invention.

Temperature control within turbine engines is highly important in order to reconcile the conflicting desire for high temperature operation for engine efficiency, whilst ensuring the engine components remain within their physical capabilities. As indicated above, generally air is drawn from the compressor side of the engine in order to cool components in the hotter turbine side of the engine, but there are still limitations upon the potential for such cooling. In such circumstances, the present invention relates to providing heat shielding across the rotor blade platforms in order to protect by limiting heat transfer to a rotor disc upon which the blades are mounted. Fig. 2 is a schematic illustration of a rotor blade arrangement 100 in accordance with the present invention. The arrangement comprises an assembly of two rotor blades 101, 102 in juxtaposed position with blade platform segments 103, 104 arranged such that their edges are in a juxtaposed joint position. A damper member 106 is located such that a rib 107 which extends radially outwards from a lateral base of the damper member 106 is in engagement with the platform 103, 104, normally as shown about the joint 105. In such circumstances the damper member 106 facilitates resilient retention of the blades 101, 102 with limited vibration chatter and good gas seal. The blades 101, 102 are secured into a rotor disc 108 through root segments 109, 110 which as shown are in the form of so called fir tree roots.

As indicated above typically a coolant air passage network will also be provided within the rotor disc 8 and blades 101, 102, but this is not shown for clarity. It will be understood that the area 111 about the blades 101, 102 in operation will be subjected to high temperature gases, and it is limitation of the heating effect of these gases upon the rotor disc 108 which is of particular concern with respect to the present invention.

The present invention provides within the damper member 106 appropriate heat shielding features in order to inhibit heat transfer from particularly the blade platform sections 103, 104 to the rotor disc 108. At a primary level the laterally extending base segments of the damper member 106 simply provide a physical barrier to convection heat flow towards the rotor disc 108 and so limit heat transfer by that mechanism. In order to enhance that barrier effect the damper member 106 may be coated with an appropriate heat conductive low heat emisivity surface or the member 106 itself either be formed from or incorporate material segments which facilitate heat transfer away from the disc 108 or inhibit heat radiation and/or transmission to that disc. In any event, the heat shielding features are normally integral with the damper member 106.

It will also be understood by appropriate shaping of the damper member 106 heat transfer may be reduced to the rotor disc 108. Such shaping may include relatively increasing surface area by ribbing or through sloping as depicted in Fig. 2. In any event, the damper member 106 in accordance with the present invention provides the dual functions of damping vibration chatter/gas seal between the blades 101, 102 as well as inhibiting through a heat shielding effect transfer of heat from the platform segments 103, 104 to the disc 108. The damper member 106 provides a radiation shield as well as the physical barrier described above between the underside of the turbine blade platforms 103, 104 and the top surface portions 112 of the rotor disc 108. This heat shield feature as indicated is most conveniently provided as an integral part with the damper member 106 and may be supported by other damper features within a cavity 113.

Generally the platform segments 103, 104 of the rotor blades 102, 103 will be able to tolerate peak temperatures in the order of 1,100 C, that is at the location A depicted in Fig. 2. Normally the rotor disc 108 must have an operational temperature which should be around 600 C for operational capability. In such circumstances typically the damper member 106 in accordance with the present invention will have a temperature in the order of 920 C.

It will be understood that heat transfer by radiation is generally to the fourth power of temperature and so by judicious formation and positioning of the damper member 106, it is possible to achieve a heat flux reduction in the order of 50% across the arrangement 100 from the platforms 103, 104 to the upper surface 112 of the rotor disc 108.

By reducing the heat flux it will be understood that other cooling mechanisms, that is to say convection and conduction, will then be able to more adequately control rotor disc 108 temperature. In such circumstances, the use of heat shielding features in the damper member 106 will contribute to cooling efficiency and will generally achieve in the order of 10 to 20 C extra relative cooling.

It will be understood that typically coolant air will be directed towards the damper member 106 including the heat shielding feature such that the effective radiation shield formed within the damper member 106 will be cooler than the radiation balance temperature. This lower temperature will be due to the cooling air in the cavity 113 being below 900 C, so in turn giving less heat radiation flux directed towards heating the rotor disc 108, and in particular the upper surface 112.

Fig. 3 illustrates a bottom perspective view of a rotor blade arrangement 200 in accordance with the present invention. Generally, a heat shield 201 is integral with and supported by an inter platform damper member (not shown) and will normally be bathed in cooling air from an appropriate source. It will be noted that the space between rotor blade shanks 202, 203 is irregular such that the lateral base portions or wings of the damper member upon which the shield 201 is formed must coincide with that irregular shaping to the space between the shanks 202, 203.

The base upon which the shield 201 is formed will be arranged to be balanced under operational loads along a centre line X-X of the damper member such that it will not interfere with damping function as described previously with respect to vibration chatter. In such circumstances there will be no contact with the blades except for the desired contact upon inclined surfaces of the damper member for appropriate function.

Considering both Figs. 2 and 3 it will be noted that the damper member 106 (Fig. 2) incorporating a radiation shield 201 (Fig. 3) is formed integrally with base portions or wings 204, 205 arranged to shield the upper surface 112 from high temperatures incident at the blade platforms 103, 104. Damping function is unaffected as the base or wing portions 204, 205 are not in contact with the blade shanks 202, 203. In such circumstances the damper member 106 behaves in a manner consistent with known damping configurations with contact between the rib 107 and the platform sections 103, 104 particularly through the joint 105. In such circumstances heat radiation flux is reduced as well as a physical barrier presented in order to inhibit heat transfer from the platform sections 103, 104 to the rotor disc 108.

It will be understood that particularly with turbine engines used with respect to aircraft that weight is important in addition to achieving satisfactory dynamic behaviour for the blades 101, 102 and the rotor disc 108.

In short, it is important that the assembly is in rotational balance. It will be understood that a number of blades will be secured around the circumference of a rotor disc and therefore care must be taken with respect to weight distribution in order to achieve such satisfactory dynamic behaviour of the assembly or blades and disc. In order to achieve this weight distribution, as illustrated in the schematic cross-section of Fig. 4 a hollow cavity 300 may be provided within a damper member 301 in order to reduce the weight of the damper 301. Furthermore, when attempting to balance rotation of a rotor blade assembly, it will be appreciated that this cavity 300 may be utilised for accommodation of balancing weights as required or necessary. The cavity 300 also provides a further means for cooling of the damping member 301 thereby further reducing its temperature and potential for radiation heat flux presentation to the rotor disc is reduced.

As indicated above, generally a large number of rotor blades will be secured about a rotor disc in order to provide a turbine stage within a turbine engine. A damper member in accordance with the present invention will be located between adjacent rotor blades in the cavity formed below the rotor blade platforms. In such circumstances the heat shielding features within the damping members will reduce the heating effect upon the mounting rotor disc and so reduce its operational temperature. Clearly, the number of blades and their particular orientation and manner of fixing to the rotor disc will depend upon operational requirements. The assembly of rotor blade arrangements between adjacent rotor blades in accordance with the present invention as indicated will form a turbine stage such as 16, 17, 18 (Fig. 1) in a turbine engine, and as indicated will add to the cooling effect upon the rotor disc by allowing such other mechanisms as convection and conduction to more effectively operate within that engine.

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 turbine blade arrangement comprising a rotor disc upon which rotor blades are secured, a damper member between root sections of respective adjacent rotor blades and the damper member having a heat shield feature to inhibit heat transfer towards the rotor disc through blade platform sections extending between the adjacent rotor blades.

2. An arrangement as claimed in claim 1 wherein the heat shield feature is integral with the damper member.

3. An arrangement as claimed in claim 1 or claim 2 wherein the heat shield feature comprises a preferential heat conduction element within the damper.

4. An arrangement as claimed in any of claims 1, 2 or 3 wherein the heat shield feature comprises appropriate heat reflective or low emissivity coatings upon the damper member.

5. An arrangement as claimed in any preceding claim wherein the heat shield feature may comprise appropriate shaping of the damper member to facilitate heat dissipation.

6. An arrangement as claimed in any preceding claim wherein the damper and heat shield feature are particularly configured for dynamic balance within a turbine blade arrangement for appropriate performance when that turbine blade arrangement is rotated.

7. An arrangement as claimed in claim 6 wherein such configuration of the damper and heat shield feature comprises a hollow section within the damper member appropriately shaped for dynamic balance within the turbine blade arrangement.

8. An arrangement as claimed in any preceding claim wherein the damper member comprises a base section arranged to extend between root sections of the adjacent rotor blades and a radial rib upstanding from the lateral base section to engage the blade platforms.

9. An arrangement as claimed in claim 8 wherein the rib engages at a juxtaposed joint between adjacent platform sections.

10. An arrangement as claimed in claim 8 or claim 9 wherein the radial rib comprises a plurality of spikes or a ridge extending along the lateral base member.

11. A turbine blade arrangement substantially as hereinbefore described with reference to the accompanying drawings.

12. A turbine blade assembly comprising a plurality of turbine blade arrangements as claimed in any preceding claim.

13. An assembly as claimed in claim 11 wherein that assembly incorporates coolant flow means and that coolant flow means is arranged to be directed in use towards the damper member in order to facilitate the heat shield feature.

14. A damper member for a turbine blade arrangement, the damper member comprising a heat shield feature to inhibit in use heat transfer across a turbine blade arrangement.

15. A member as claimed in claim 14 wherein the heat shield feature comprises a low emissivity coating or element or shaping of the damper member.

16. A member as claimed in either claim 14 or claim 15 wherein the number has a hollow passage or cavity to provide in use dynamic balance for the damper member.

17. A turbine engine incorporating a turbine blade arrangement as claimed in any of claims 1 to 12 or a turbine blade assembly as claimed in claim 12 or claim 13 or damper member as claimed in claim 14 to 16.

18. 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.