GB2084654A

GB2084654A - Cooling gas turbine engines

Info

Publication number: GB2084654A
Application number: GB8126956A
Authority: GB
Original assignee: MTU Motoren und Turbinen Union Muenchen GmbH
Current assignee: MTU Aero Engines GmbH
Priority date: 1980-10-01
Filing date: 1981-09-07
Publication date: 1982-04-15
Also published as: GB2084654B; DE3037020A1; FR2491142A1; FR2491142B1; JPS5781129A; DE3037020C2

Abstract

Cooling air is tapped from the engine compressor 1 and is ducted to the turbine 3 via a duct 9 formed between rotationally symmetrical inner and outer wall sections 4, 5, 6 and 7, 8 respectively of the common compressor-turbine rotor system and in which guide plates 13, 14 are arranged to deflect the cooling airflow. Cooling is thereby possible with greatly reduced losses. <IMAGE>

Description

SPECIFICATION Improvements in and relating to gas turbine engines This invention relates to a gas turbine engine, more particularly a multiple-shaft turbojet engine, in which air is tapped from the compressor, especially of the high-pressure compressor, and is ducked to the turbine, preferably the high-pressure turbine for cooling purposes.

In order to achieve adequate cooling, especially in the high pressure turbine of large modern turbojet engines, cooling air under a relatively high pressure, is blown from the leading edges of the rotor blades, sometimes against the direction of turbine flow. This air is generally diverted from the combustion chamber area (at a point between the outer casing and the flame tube of the combustion chamber) and led, through preswirl nozzles, to the rotor or, more specifically, to the turbine wheel disk. This transition from the stationary to the rotating system involves a number of losses: for example, mixing losses, shock losses, and especially leadage losses due to the requisite high cooling air pressure.

According to this invention we propose a gas turbine engine wherein cooling air is tapped from the compressor, and passes through a ducttothetur- bine for cooling purposes, wherein the said duct forms part of a common compressor-turbine rotor system. Other features of the invention are set forth in the appendant claims.

The invention enables turbine cooling without, or at least with much reduced, losses as compared with conventional arrangements. Also, it is possible to achieve aerodynamically favourable flow conditions while ensuring a relatively high cooling air pressure for turbine cooling. Comparatively little power need be extracted from the aero-thermodynamic cycle process of the engine.

One embodiment of the present invention will now be described by way of example with reference to the accompanying drawings which is an axial cross section of part of a gas generator and associated high-pressure rotor system of a multiple-shaft turbojet engine.

The gas generator of the turbojet engine illustrated in the drawing comprises a mutiple-stage axial-flow high-pressure compressor 1, an annular combustion chamber 2 and a single-stage, axial-flow high-pressure turbine 3. The high-pressure compressor 1 and the high-pressure turbine 3 (turbine disk 10) are interconnected by means, the highpressure rotor spool which is made up of rotationally symmetrical or barrel-shaped radially inner4, 5 and 6 and outer 7,8 wall sections defining there between a duct 9 through which cooling air tapped from the compressor flows to the turbine wheel 10 and its turbine rotor blades 11. This makes the air tapping and ducting provisions in the high-pressure rotor operationally co-rotating.

The wall sections 4 to 8 which make up the duct 9 are joined together as shown at A and B so separating the duct into regions. Circumferential slots or the like are provided for the flow of air from one region to the other, the slots at A, being offset, that is angularly displaced, relative to the slots at B.

In the illustrated embodiment, the cooling air is diverted from the final stage of the high-pressure compressor 1, through a bleed slot 12 behind the rotor blades 11 and coaxial with the centreline of the engine. The slot 12 is defined betweed the peripheral edge of the wheel disk face and a lower leading-edge portion of the stator of the final compressor stage.

Depending radially inwardly from the compressor stator at the entrance of the duct 9, is a wall section 12' which serves to provide an aerodynamically favourable air bleed.

Within the duct 9 are first 13 and a second 14 rings of blades the first ring 13 being mounted on the wall section 4 in a curved section of the duct 9 deflecting the cooling air flow from the radial to the axial direction and the second ring 14 being mounted on a radially extending portion section 8 of the wall immediately upstream of the turbine wheel disk 10, to which wall section 8 is connected.

The tips of the blades in the first ring 13 are spaced at a moderate distance from the radially outer wall section 7, and the tips of the blades in the second ring 14 are axially spaced from the adjacent turbine wheel disk 10, along a major part of their length.

These blade tip spacings are intended to accom modatedifferential expansion of the various components.

The portion 8' of the wall section 89 adjacent and connected to the turbine disk 10 exhibits an end portion 16 curved toward the wheel disk 10 and intended to achieve locally selective, flow-promoting guidance of the cooling air flow to the air-cooled turbine rotor blades 11 via recesses 17,18, Aj8t 19 in the wheel disk. Spent cooling air expelled from the blades and into the gas flow in the turbine duct 20 through outlet slots or holes arranged in the leading edge of each blade. The cross-section of the duct 9 in the region of the second ring of blades 14 on the turbine side initially converges in the direction of cooling airflow and then diverges to produce a cooling air velocity sufficient for the intended cooling effect.

In operation, cooling air is diverted from the high-pressure compressor 1 - at a sufficiently high pressure for high-pressure turbine rotor blade cooling through the bleed slot 12 between the rotor and stator blades 11 and 11', respectively, and is ducted radially inwards, where the circumferential component of the flow, increases in a direction inward from the rotor wheel outlet in accordance with the vortex distribution (approximately represented by the free angle law) until the flow assisted by wall friction in the duct 9 adapts to the circumferential speed of the rotor. The blade ring 13 begins at the radius VE at which the circumferential velocity component of the flow reaches the circumferential velocity of the rotor.

From this point on until the flow reaches the minimum radius Vmin of the flow duct the blades 13 guide the flow and extracts from it energy (turbine action). Over the radial-component distance on the turbine side, the circumferential speed of the highpressure rotor is imparted to the cooling air via the second ring of blades portion 14 to minimize wall friction losses and ensure shock-free entry of the cooling air into the high-pressure turbine rotor blades 11. Energy, therefore, is supplied to the flow through the compressor action of the blade ring 14 from the inlet radius VT1 thereof to the outlet VT2.

Claims

CLAIMS:

1. . A gas turbine engine wherein cooling air is tapped from the compressor, and passes through a duct to the turbine for cooling purposes, wherein the said duct forms part of a common compressorturbine rotor system.

2. An engine according to Claim 1, wherein the duct is formed between rotationally symmetrical inner and outer wall sections of the common rotor system.

3. An engine according to Claim 2, wherein an outer wall section sot the entrance of said duct is arranged on the stator of the compressor.

4. An engine according to any one of claims 1 to 3 comprising a first set of blades first arranged in curved section of the duct for deflecting the cooling air flow from the radial to the axial direction, and a second set of blades arranged in a radially extending section upstream of and connected to the turbine wheel disk of the common rotor system.

5. An engine according to Claim 4, wherein the blades of the first set of blades are mounted on the inner wall section of the duct, the tips of the said blades being spaced from the outer wall section of the duct.

6. An engine according to claim 4 or claim 5, wherein the blades in the second set of blades are attached to an outer wall section of the duct adjacent the turbine wheel disc with the tips of the blades, along at least a part oftheir length, axailly spaced from the turbine wheel disc.

7. An engine according to any one of Claims 1 to 6, wherein a wall section of the duct adjacent and connected to the turbine wheel disc is shaped to deflect the cooling airflow via the wheel disc to the turbine rotor blades.

8. An engine according to any of Claims 1 to 7, wherein in the region of the second set of blades the duct initially converges in the direction of cooling air flow and subsequently diverges.

9. A gas turbine engine constructed and arranged substantially as hereinbefore described with reference to and as illustrated in the accompanying drawing.