US2912221A

US2912221A - Apparatus for cooling turbine wheels in combustion turbines

Info

Publication number: US2912221A
Application number: US467576A
Authority: US
Inventors: Chamberlin Reginald He Douglas
Original assignee: D Napier and Son Ltd
Current assignee: Napier Turbochargers Ltd
Priority date: 1953-11-20
Filing date: 1954-11-08
Publication date: 1959-11-10
Anticipated expiration: 1976-11-10
Also published as: BE533517A; GB789203A; CH341029A; FR1114538A; NL192509A; NL99245C; DE1023275B

Description

Nov. 10, 1959 R. H. D. CHAMBERLIN v 2,912,221

APPARATUS FOR COOLiNG TURBINE WHEELS IN COMBUSTION TURBINES Filed Nov. 8, 1954 4 Sheets-Sheet 1 Inv entor $7 4M #1QW M Attorney Nov; 10, 1959 Filed Nov. 8, 1954 R. H. D. CHAMBERLIN APPARATUS FOR COOLING TURBINE WHEELS IN COMBUSTION TURBINES 4 Sheets-Sheet 2 Attorney5= Nov. 10, 1959 R. 1-1. D. CHAMBERLIN 2,912,221

APPARATUS FOR COOLING TURBINE WHEELS IN COMBUSTION TURBINES Filed Nov. 8, 1954 4 Sheets-Sheet 3 wwmwwzwm WW 0 5 55525 4555553 Inventor Nov. 10, 1959 R. H. D. CHAMBERLIN 2,912,221

APPARATUS FOR COOLING TURBINE WHEELS IN COMBUSTION TURBINES Filed Nov. 8, 1954 4 Sheets-Sheet 4 Inventor 0 444. 0. W

M I- u/at'ln APPARATUS FOR COOLING TURBINE WHEELS IN COMBUSTION TURBINES Reginald Henry Douglas 'Chamberlin, Ealing, London,

England, assignor to D. Napier & Son Limited, London, England Application November 8, 1954, Serial No. 467,576

Claims priority, application Great Britain November 20, 1953 Claims. (Cl. 253-39115) This invention relates to combustion turbines, that is to say, to turbines driven by the hot gaseous products of combustion and is particularly though not exclusively applicableto gas turbine power plants as used for aircraft propulsion purposes.

It is an object of the invention to provide improved cooling for such combustion turbines.

According to the invention, a combustion turbine of the kind referred to, includes a turbine rotor disc carrying a ring of rotor blades and a ring of stator blades lying upstream, and/or downstream of the rotor ring, a stationary annular body arranged to lie adjacent to one face of the rotor disc, means for admitting cooling fluid to the inner circumferential parts of the annular space between the rotor disc and the annular body, the annular body being formed or provided on its face which lies adjacent to the face of the rotor disc with a series of spiral fins, such that the cooling fluid tends to How in a spiral path over the rotor disc during its passage outwards between the rotor disc and the stationary annular body.

The invention is particularly applicable to turbines which include two or more rotor discs. Thus according to another aspect of the invention a combustion turbine includes two rotor discs each carrying a ring of rotor blades, and a ring of stator blades lying between the two adjacent rotor blade rings, an annular body connected to and extending inwards from the stator blade ring so as to lie between but spaced from the adjacent faces of the rotor discs, and means for admitting cooling fluid to the inner circumferential parts of the spaces between the rotor discs and the intermediate annular body, the annular body being formed or provided on each of its faces which lies adjacent to the. face of a rotor disc with a series of spiral fins, such that the cooling fluid tends to flow in a spiral path over the rotor discs during its passage outwards between the rotor discs and the annular body.

Preferably the annular body is in the form of a double-walled hollow annular shell, the opposite side walls of which carry on their external surfaces the spiral fins, and are stiffened internally by spacer plates extending between them.

Moreover the annular body is preferably of such form and dimensions and is so arranged that its side surface or surfaces lie closely adjacent to the face or faces of the rotor disc or discs as the case may be.

According to another preferred feature of the invention the axial clearance between the annular body and the adjacent face of a rotor disc, is approximately $5 of the axial clearance between adjacent rotor discs.

The turbine preferably includes a sealing member mounted on the inner part of the annular body, and engaging with a rotary sealing member mounted on the rotor shaft, so as to provide an interstage fluid seal between the upstream and downstream sides of the stator blade ring, the radial displacement of the sealing mem- Patented Nov.,10, 1959 2 I her from the axis of the rotor shaft being considerably less than the radial displacement of the stator blade roots from this axis.

In such case the means for admitting cooling fluid to the spaces between the rotor discs and the annular body are preferably such as to create a dilferential pressure acting on the rotor discs in an axial direction to oppose the dynamic thrust caused by the gas flow through the turbine.

According to another aspect of the inventiona gas turbine power plant includes a combustion turbine arranged as referred to above, and a main compressor mounted coaxially with and driven by the turbine, and arranged to deliver compressed air to one or more combustion chambers from which the products of combustion are led to the turbine, and the cooling fluid admitted to the annular space or spaces between the rotor disc or discs and the annular body is derived from the main compressorand is admitted to the said annular space or spaces via one or more inwardly directed passages in the compressor rotor, and a conduit in the compressor and turbine rotor shafts.

The power plant preferably includes a rotary thrust balance piston mounted on the rear end of the turbine shaft, and means'for admitting air under pressure to the piston from the conduit in the turbine rotor shaft.

The invention may be performed in various different ways but one specific embodiment will now be described by way of example, as applied to the exhaust gas turbine of an aircraft propeller-turbine power installation, and with reference to the accompanying drawings, in which:

Figure 1 is a side elevation, partly broken away and partly in section, of the complete power unit,

Figure 2 is an enlarged half sectional view of the exhaust turbine,

Figure 3 is a half sectional view of the turbine, on a further enlarged scale, and

Figure 4 is an end elevation of one of the annular bodies 28 partly-broken away to show a cross section on the line 4-4 in Figure 2.

The power unit comprises basically an axial flow compressor 1, arranged to deliver compressed air to a series of combustion chambers 2 Where the air is burnt in known manner with fuel supplied through supply passages 3, the products of combustion being fed to the working plenum of a turbine 4, and then passed to exhaust through the exhaust nozzle 5. The shaft 6 of the turbine 4 is directly coupled to the shaft 7 of the compressor 1, by means of a hollow tubular assembly 8 and the forward end of the compressor shaft is connected through a differential gear indicated generally at 9 in Figure 1 to a shaft 10 on which is mounted a propeller (not illustrated).

The turbine comprises three

rotor discs

21, 22, 23

mounted for rotation on the rotor shaft 6, each rotor disc carrying at its outer periphery a series of spaced turbine rotor blades 25, constituting three turbine rotor rings.

interposed between the rotor rings and upstream thereof, are corresponding stator rings 26 each comprising aseries of circumferentially spaced stator blades With their outer ends anchored to a stator casing 27.

The inner ends of the blades 26 of each of the intersta ge stator rings are connected to hollow annular interstage bodies 28, 29 formed of sheet metal and lying between adjacent rotor discs '21, 22, 23 and spaced therefrom by a small axial clearance. Each group of three adjacent stator blades are mounted at'their inner ends on a common platform 30 (see Figure 4) which is pivotally connected to a point on the periphery of the respec tive body 28, or 29, by means of a tangential link 31, thus permitting thermal expansion of the blades in a radial direction, while yet providing accurate and rigid support for the interstage bodies 28, 29, concentric with the turbine shaft 6. In the present example where the axial distance between

adjacent rotor discs

21, 22, 23 is approximately 2 inches the axial clearance between a rotor disc and the adjacent face of the adjacent interstage body 28, 29, is approximately 0.2 inch. The inner cylindrical surface of each interstage body is provided with a labyrinth seal 32 engaging a cylindrical fiange formation 33 on a sealing member 56 arranged between each pair of adjacent rotor discs. Each sealing member is provided with radial drillings, passages or

cutaway portions

34, 35, which communicate with radial drillings as in the rotor shaft to permit cooling air to flow outwards from the hollow rotor shaft into the clearance spaces between each interstage body 28, 29, and the

adjacent rotor discs

21, 22, 23.

The cooling air is derived from the main compressor 1, and is fed from immediately upstream of the last stage 38 thereof inwards into a conduit 37 within the compressor rotor shaft, assisted by vanes provided on one of the compressor rotor discs and arranged to act as a centripetal pump. The conduit 37 communicates directly with the interior of the turbine rotor shaft 6 through the hollow tubular connecting assembly 8.

The drillings 3d admitting cooling air to the annular spaces on the downstream sides of the rotor discs are larger than the drillings 35 admitting air to the upstream sides of the discs and thus tend to produce higher air pressures on the downstream sides of the discs so as to provide a counterbalancing axial thrust on the rotor of the turbine in a direction to resist the dynamic thrust caused by the gas fiow through the turbine. The outlet of cooling air from the periphery of each annular space referred to is restricted by the usual knife edge seals 39 formed on the roots of the rotor blades 25.

Each interstage body 28, 29, is formed from two spaced annular sheet metal discs 4t), 31, having inturned inclined flanges at their inner and outer peripheries, the extreme inner and outer edges of the

discs

40, 41 being welded respectively to an inner annular ring 42 carrying the labyrinth seal 32, and an outer ring 43 to Which the inner ends of the stator blades 26 are connected as described above.

The interior of each interstage body is vented through a drilling 53 in the mounters 42 and 32, to a midpoint in the labyrinth seals 32, 3 3, thus maintaining an intermediate pressure in the interiors of the bodies, and reducing the differential pressure on each

wall

40, 41, which can therefore be of relatively light construction.

Each interstage body 28, 29 is braced internally by a series of radial spacer plates 44 of approximately rectangular shape and having a number of cutaway apertures to reduce weight and to permit air to circulate within the body. Although anchored only at its outer periphery each body is thus sufficiently stiff to withstand the combined pressure differences acting on its opposite external annular surfaces.

The external upstream and downstream faces of the

discs

40, 41 of each of the interstage bodies are provided with spiral fins do which may be of sheet metal welded to the body, and which tend to cause the cooling air to circulate outwards in spiral fashion over the face of the

rotor discs

21, 22, 23, and thus to increase the length of the path and hence the cooling effect of the air in relation to the quantity of air passing.

Preferably the spiral fins 46 are arranged as shown in Figure 4 to provide relatively narrow spiral passages 47 which have a circumferential component of direction considerably greater than the radial component. Thus in the preferred illustrated construction five fins are provided, each extending round an angle of approximately 300, the inner end of each fin bein approximately 3.75 inches from the axis of he turbine rotor, while the outer end is approximately 6.5 inches from this axis. In this 4 case the width of the passages 4'7 between the fins is approximately 0.4 inch. The fins are equally radially spaced relative to the axis of the turbine and each fin subtends an angle at the turbine axis which is substantially greater than the angle subtended by the inner ends of a pair of adjacent fins.

Corresponding spiral fins 46 are also provided on the stator Walls 48 lying adjacent the upstream surface of the first rotor disc 21 and the downstream surface of the last rotor disc 23, cooling air being supplied also to the clearance spaces between these walls and the adjacent rotor discs by way of port 57 from space 54 which is in communication with the outlet of the compressor, and from air supply passage 55 respectively.

On the rear end of the hollow turbine shaft 5 is mounted an additional thrust balancing piston 56, formed with a labyrinth seal 51 at its periphery which cooperates with a fixed part of the turbine casing to provide a substantially air tight seal therewith. The piston 5% is enclosed within a domed cover 52, also secured in gas tight manner to the turbine casing, and air from the hollow interior of the turbine shaft, 6, which is under substantial pressure is admitted direct to the space between the piston and the end cover, and thus exerts a counterbalancing axial thrust on the piston 56 and turbine shaft in an upstream direction to assist in counteracting the dynamic thrust in an axial direction caused by the flow of gas through the turbine.

The diameter of the rotor rings at the roots of the rotor blades 25 is in the present example 16.6 inches. The diameter of the first sealing member 33 between the first and second rotor discs 2f, 22, where it engages the labyrinth seals 32 on the inner surface of the first interstage body 28, is 5.4 inches, while the corresponding diameter of the second sealing member 33 is 5.0 inches. The relative speed of the two contacting parts of each of these labyrinth seals varies of course as a direct multiple of their diameter, and in the construction described accordingly provides relative velocities of the order of a third of the velocity which would obtain if the interstage seals were adjacent the roots of the blades 25 in the rotor rings. Moreover the construction provides at the same time efficient cooling of the rotor discs blades and associated parts, and a certain degree of axial balancing thrust, and is light and compact.

What I claim as my invention and desire to secure by Letters Patent is:

l. A combustion turbine comprising a hollow turbine rotor shaft, means operatively connected to and supplying cooling fluid under pressure into said hollow shaft, two rotor discs mounted for rotation with the shaft and a sealing member secured between said rotor discs and having passages radially therethrough, a ring of rotor blades secured to each rotor disc, a stator casing surrounding the working plenum of the turbine and a ring of stator blades secured at their outer ends to the stator casing and constituting a stator blade ring lying between the two adjacent rotor blade rings, an annular body connected to and extending inward from the stator blade ring so as to lie between the adjacent rotor discs and adjacent to said discs although spaced from said rotor discs to define annular spaces therebetween, said shaft being formed with drillings extending from its hollow interior and communicating with the annular spaces between said rotor discs, and the annular body through said passages, for admitting cooling fluid to said spaces, and a series of spiral fins formed on one of the faces of the annular body which lies adjacent one of the respective rotor discs, such that the cooling fluid tends to flow in a series of spiral paths over one of the rotor discs during its passage outwardly through the annular spaces.

2. A combustion turbine as claimed in claim 1, in which the annular body is in the form of a double-walled hollow annular shell, and comprises two annular side walls which carry on their external surfaces the spiral fins, and a series of internal spacer plates extending between the said side walls.

3. A combustion turbine as claimed in claim 1, in which the side surfaces of the annular body lie closely adjacent to the faces of the rotor discs on either side thereof.

4. A combustion turbine as claimed in claim 3, in

which the clearance between the side walls of the annular body and the adjacent rotor discs is approximately one tenth of the axial distance between adjacent rotor discs.

5. A combustion turbine as claimed in claim 1, in which each fin subtends an angle of approximately 300 at the turbine axis.

References Cited in the file ofthis patent UNITED STATES PATENTS Whittle u. Ian. 6, 1948 Johnstone May 27, 1952 Morley et a1 Dec. 2, 1952 McLeod Nov. 3, 1953 Lundquist Mar. 16, 1954 FOREIGN PATENTS France May 12, 1947