GB2049068A - Turbine bladed rotors - Google Patents

Abstract

A turbine 10 suitable for a gas turbine propulsion engine comprises a rotor disc 12 provided with an annular array of blades 14 which may be formed from silicon nitride. Each space between the radially inner regions of adjacent blades 14 is occupied by a refractory insulating member 19. The refractory members 19 together protect the disc 12 from the thermal effects of the hot gas stream which passes in operation through the turbine 10 as well as providing a certain degree of blade damping. The refractory insulating members 19 are formed from a low density fibrous refractory material such as fibrous alumina bound by an organic adhesive or colloidal oxide provided with a densified skin to ensure erosion resistance. <IMAGE>

Description

SPECIFICATION Fibrous alumina turbine insulation block This invention relates to turbines and in particular to turbines suitable for gas turbine propulsion engines.

The turbines of gas turbine propulsion engines conventionally comprise axially alternate annular arrays of rotary and stator aerofoil blades. Each array of rotary blades is mounted on a disc and each blade in the array provided with two circumferentially extending platforms. The platforms of adjacent blades cooperate to define a portion of the radially inner wall of the annular passage through which flow in operation the gases passing through the turbine. The radially outer wall of the annular passage is defined by the casing of the turbine.

The gases which pass in operation through the turbine are usually extremeiy hot and consequently the blade aerofoil portions and platforms are rapidly heated up to very high temperatures, Since turbine blades are conventionally formed from thermally conductive materials, they in turn heat up the disc on which they are mounted by conduction through the blade roots and by convection and radiation from those blade regions, such as the blade platforms, which are spaced apart from the disc. The net result is that under normal operating conditions, the disc has a very high running temperature. This in turn dictates the use of expensive high temperature alloys in disc manufacture and perhaps additionally the use of cooling air derived from the gas turbine engine compressor in order to reduce disc temperatures, especially in the region of the rim of the disc.

It is an object of the present invention to provide a turbine suitable for a gas turbine propulsion engine which is so arranged that the amount of heat transferred from the hot gas stream passing operation through the turbine to the discs on which the rotary turbine aerofoil blades are mounted is reduced According to the present invention a turbine suitable for a gas turbine propulsion engine comprises radially inner and outer walis defining an annular passage adapted to contain the motive fluid passing in operation through said turbine and at least one rotary disc mounted coaxially with said annular passage and carrying an annular array of spaced apart aerofoil blades, each blade extending transversely across said annular passage, each space between the radially inner regions of adjacent aerofoil blades being occupied by a refractory insulating member, said members together being adapted to define a portion of the radially inner wall of said annular passage and provide a thermal barrier between said motive fluid and said disc, means being provided to retain each refractory insulating member in position in its corresponding space between the radially inner regions of said adjacent aerofoil blades.

Said refractory insulating members are preferably retained radially against centrifugal loading by engagement means provided on said radially inner regions of said aerofoil blades.

Said engagement means may comprise angled portions so arranged that said radially inner regions of adjacent aerofoil blades define radially outwardly converging seatings for said refractory insulating members.

Alternatively said engagement means may comprise circumferentially extending portions adapted to engage the radially outer regions of said refractory insulating members.

Each of said refractory insulating members is preferably formed from a low density fibrous refractory material and is provided with a densified skin.

Said low density fibrous refractory material may comprise fibrous alumina and a binder.

Alternatively said low density fibrous refractory material may comprise fibrous zirconia and a binder.

Said binder may be aniorganic adhesive or a colloidal oxide.

Said densified skin on each of said refractory insulating members may be formed by plasma arc spraying the surface of said members with a high melting point refractory material.

Said high melting point refractory material may be alumina, zirconia, a silicate or a zirconate.

Said aerofoil blades may be formed from a ceramic material.

Said aerofoil blades may be formed from silicon nitride.

The invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 is a sectioned side view of a portion of a turbine in accordance with the present invention.

Figure 2 is a sectioned side view of an alternative form of construction of the turbine shown in Figure 1.

Figure 3 is a.yiew on A-A of the turbine portions shown in Figure 1.

Figure 4 is a view on A-A of Figure 1 of an alternatlve.form of construction of the turbine shown in Figures 1 and 3.

With reference to Figure 1, a turbine 10 suitable for a gas turbine propulsion engine comprises a casing 11 enclosing a rotary disc 12. The disc 12 is mounted coaxially within the casing 11 so that an annular passage 1 3 is defined between them.

The annular passage 1 3 is adapted to contain the gases passing in operation through the turbine 10.

A plurality of similar aerofoil blades 14 are equally spaced around the periphery of the disc 12 so as to extend transversely across the annular passage 13 and are anchored to the disc 12 by means of part-circular cross-section roots 1 5. In addition to the roots 15, each aerofoil blade 14 is provided with a shank portion 1 6 and an aerofoil portion 17, the shank portion 1 6 interconnecting the root portion 1 6 with the aerofoil portion 1 7.

The shank portion 1 6 of each aerofoil blade 14 is provided with diverging faces 1 8 so that the faces 1 8 of adjacent blades converge in a radially outward direction. The faces 1 8 provide seatings for refractory insulating members 1 9 which are positioned between adjacent aerofoil blades 14.

Thus the radially outward convergent configuration of the faces 1 8 of adjacent aerofoil blades 14 ensures that upon rotation of the disc 12, each refractory insulating member 1 9 is restrained radially against centrifugal loadings.

In the alternative form of turbine construction shown in Figure 2, each refractory insulating member 1 9 is restrained radially by means of circumferentially extending steps 20 provided on each blade shank portion 16. The steps 20 engage the radially outer regions 21 of each refractory insulating member 19 sofas to provide such radial restraint. It will be appreciated, however, that alternative shank and refractory insulating member configurations could be employed for the radial restraint of the refractory insulation members 19.

The refractory insulating members 19 are restrained axially by means of side plates 22 which are located one each side of the disc 12 and extend across the ends of the refractory insulating members 1 9 as can be seen in Figure 3. Other forms of axial restraint could however be employed. Thus for instance in the embodiment shown in Figure 4, the disc 12 is provided with flanges 23 which extend across the ends of the refractory insulating members 19. In order to provide access to the refractory insulating members 9 the disc 12 is radially split into two separable portions 24 and 25.

The radially outer surfaces 26 of the refractory insulation members 19 9 bridge the gaps between the radially inner regions of the aerofoil blades 14 so as to define a portion of the radially inner wall of the annular'passage 13. Consequently the refractory insulation members 14 provide a thermal barrier between the disc 12 and the hot gases which pass in operation through the annular passage 13. This means that cooling of the disc 12 is unnecessary and indeed that the disc 12 may be manufactured from cheaper, lower temperature alloys than those conventionally used.

The refractory insulation members 1 9 are formed from a low density fibrous refractory material in order to ensure that the centrifugal loadings they impose upon the aerofoil blades 14 are as low as possible. Thus the refractory insulation members 19 may be formed from fibrous alumina or zirconia bound together with a suitable organic adhesive or colloidal oxide. Such materials usually have densities in the range 0.35 to 1.0 grams per cc. Additional advantages of such fibrous refractory materials are that as well as fixing the aerofoil blades 14 in their correct upright position, they also provide a certain degree of blade damping, and are resistant.to thermal shock and oxidation.

It is possible that during turbine operation, the refractory insulation members 1 9 will be subjected to erosion. In order to prevent this, we prefer to provide the members 1 9 with a densified skin. This may be achieved by plasma arc spraying the members 19 with a high melting point refractory material. Thus the members 19 may be plasma arc sprayed with alumina, zirconia, a high melting point silicate such as Mg2SiO4 or a high melting point zirconate such as MgO ZrO2 or Car04.

When turbines are required to operate at very high temperatures and high rotational speeds, it is frequently desirable to manufacture the aerofoil blades 14 from a ceramic material such as silicon nitride. However such materials are not well suited; to conventional aerofoil blade designs with their integral circumferentially extending platforms.

Such platforms give rise to high bending stresses which are unacceptable in conventional ceramics with their low tensile strengths. Since the aerofoil biases 14 of the turbine of the present invention are not provided with such platforms, it is possible to manufacture them from ceramic materials such as silicon nitride. Thus a substantial portion of the turbine 10 could be manufactured from non metallic high temperature resistant materials, thereby reducing material costs, permitting higher operating temperatures and obviating the need for air cooling of the aerofoil blades and discs.

Claims (14)

1. A turbine suitable for a gas turbine propulsion engine comprising radially inner and outer walls defining an annular passage adapted to contain the motive fluid passing in operation through said turbine and at least one rotary disc mounted coaxially with said annular passage and carrying an annular array of spaced apart aerofoil blades, each blade extending transversely across said annular passage, each space between the radially inner regions of adjacent aerofoil blades being occupied by a refractory insulating member, said members together being adapted to define a portion of the radially inner wall of said annular passage and provide a thermal barrier between said motive fluid and said disc, means being provided to retain each refractory insulating member in position in its corresponding space between the radially inner regions of said adjacent aerofoil blades.

2. A turbine as claimed in claim 1 wherein said refractory members are retained radially against centrifugal loading by engagement means provided on said radially inner regions of said aerofoil blades.

3. A turbine as claimed in claim 2 wherein said engagement means comprise angled portions so arranged that said radially inner regions of adjacent aerofoil blades define radially outwardly converging seatings for said refractory insulating members.

4. A turbine as claimeci in claim 2 wherein said engagement means comprise cir;cumferentially extending portions adapted to engage the radially outer regions of said refractory insulating members.

5. A turbine as claimed in any one preceding claim wherein each of said refractory insulating members is formed from a low density fibrous refractory material and is provided with a densified skin.

6. A turbine as claimed in claim 5 wherein said low density fibrous refractory material comprises fibrous alumina and a binder.

7. A turbine as claimed in claim 5 wherein said low density fibrous refractory material comprises fibrous zirconia and a binder.

8. A turbine as claimed in claim 6 or claim 7 wherein said binder is an organic adhesive or a colloidal oxide.

9. A turbine as claimed in any one of claims 5 to 8 wherein said densified skin on each of said refractory insulating members is formed by plasma arc spraying the surface of said members with a high melting point refractory material.

10. A turbine as claimed in claim 9 wherein said high melting point refractory material is alumina, zirconia, a silicate or a zirconate.

11. A turbine as claimed in any one preceding claim wherein said aerofoil blades are formed from a ceramic material.

12. A turbine as claimed in claim 11 wherein said ceramic material is silicon nitride.

13. A turbine substantially as hereinbefore described with reference to and as shown in the accompanying drawings.

14. A gas turbine engine provided with a turbine as claimed in any one preceding claim.