GB2170865A

GB2170865A - Gas turbine engine with cooling device

Info

Publication number: GB2170865A
Application number: GB08503286A
Authority: GB
Inventors: George Pask
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1985-02-08
Filing date: 1985-02-08
Publication date: 1986-08-13

Abstract

The turbine of a gas turbine engine which powers an aircraft experiences varying operating temperatures, e.g., at take off the operating temperature is higher than at cruise, and this necessitates the provision of a variable cooling air flow to the turbine. The invention provides a secondary cooling system comprising a diaphragm-controlled valve which has air inlet and air outlet orifices 50,52 one of two annular in a members 43,48 and is controlled by way of flexing the diaphragm 60 to allow communication therebetween, in accordance with the pressure difference across the diaphragm 60 which is varied in accordance with sensed change in turbine temperature. <IMAGE>

Description

SPECIFICATION Gas turbine engine with cooling device This invention relates to a gas turbine engine and includes a cooling device which is associated with the turbine thereof.

More specifically, the invention relates to a gas turbine engine which in operation powers an aircraft, which through its flight regime needs changing power outputs from the gas turbine engine.

According to the present invention, a gas turbine engine includes a valve device for controlling a flow of cooling air to a turbine stage comprising a chamber, one internal wall surface of which has a protuberance, a diaphragm which in one position seals against the chamber rim and the protuberance so as to define a space either side of said protuberance and a further space on the other side of the diaphragm, cooling air inlet means on the side of the protuberance, cooling air outlet means on the other side of the protuberance and connected to a turbine cooling flow path and means for enabling flexing of the diaphragm away from the protuberance and so enable communication between the air inlet means and the air outlet means, the flexing means being adapted for actuation on the onset of an operating regime which results in an increase in operating temperature of the turbine stage of the gas turbine engine.

The means for enabling flexing of the diaphragm may comprise a removable counter pressure means for, prior to actuation, countering air pressure on the air inlet side of the diaphragm so as to, in one mode of operation, maintain the diaphragm seal.

Preferably the counter pressure means comprises a pressurized air supply in communication with the further chamber and includes means for venting the air supply when the turbine stage achieves a given increase in operating temperature.

Preferably the means for venting the air supply comprises a bleed point in the air supply and a further valve blocking the bleed point and moveable to unblock the bleed point in response to a signal which is generated on achievement of the operating regime.

The further valve may comprise a solenoid device which is connected for actuation in response to the signal.

Alternatively, the further valve may be connected for actuation by a throttle control lever of the gas turbine engine, such that opening of the throttle by the lever also moves the further valve so as to unblock the bleed point.

Preferably the chamber is annular and includes a plurality of air inlets and air outlets in equi-angularly spaced relationships around the annulus.

Alternatively, an annular member may contain a plurality of the chambers, arranged in equi-angular relationships around the upstream face thereof.

Each air inlet has a cooperating air outlet which may be positioned radially inwardly thereof.

The invention will now be described, by way of example and with reference to the accompanying drawings in which: Figure 1 is a cross sectional part view of the turbine of a gas turbine engine incorporating an embodyment of the present invention.

Figure 2 is a cross sectional view of the embodiment of the present invention which is incorporated in Fig. 1.

Figure 3 is a diagrammatic arrangement of the whole of the embodiment of Fig. 1 and Figure 4 is a pictorial part view of an alternative embodiment of the present invention.

Referring to Fig. 1. A turbine 10 of a gas turbine engine (not shown) includes a stage of nozzle guide vanes 12 which is fixed between an outer casing (not shown) and frusto conical members 14, 16 and 18. A rotatable stage of turbine blades 20, mounted on a turbine disc 22 is positioned immediately downstream of the stage of nozzle guide vanes 1 2.

Frusto conical members 24 has a number of annular lands 26 on its downstream face ie., downstream with respect to the direction of flow of exhaust gases through the engine (not shown). A blank disc 28 is affixed to the turbine disc 22 for co-rotation and has corresponding lands 30 on its upstream face, which cooperate with the lands 26 to control cooling air leakage therethrough, in known manner.

The frusto conical member 18 has an annular central portion 32 to which an annular valve 34 is fixed. The valve 34 is connected to deliver a supply of cooling air via a number of equi-angularly spaced passage ways 36 in the central portion 32, to a space 38 radially inwardly of the lands 26 and 30.

A further cooling air supply is, in operation, directed via further passageways 40 in the central portion 32, which are spaced alternately with the passage ways 36, to the space 38. The further air supply does not pass through the value 34. Both air supplies however, pass through respective swirl devices 42 prior to entering the space 38. The swirl devices 42 ensure that the cooling air enters the space 38 in the direction of rotation of the disc 30, thus avoiding friction therebetween, and consequent heating of the air.

Referring now to Fig. 2. The valve 34 has a first annular member 43 the downstream face 44 of which is hollowed and then relieved so as to provide annular lands 46. The hollowing is in the form of a constant radius curve in all planes radially of the annular member 43, and provides an annular space 51. The valve 34 also has a second annular member 48 which has two annular grooves 50 and 52 machined in its upstream face, so as to form two spaces one either side of an annular land 58.

The members 43 and 48 are clamped together by peened over lips 45 on the member 43.

The hollowed spaces defined in the downstream face of the annular member 43 and the two spaces 50 and 52 are, in one operative mode, seperated from each other by an annular flexible diaphragm 60. The seperation is maintained by the upstream face of the diaphragm 60 being exposed to pressurized air in space 90 entering the upstream space 51 via a passage 100 in the annular member 43, which air pressure is also present in the space 50, by virtue of its connection via air inlet 52, with the space 90. Space 90 however, is subject to constant leakage through passage 40 and the turbine blades, whereas the space 51 is remotely sealed by a further valve 64 (Fig.

3) which is described later in this specification. Suffice it to say that the aforementioned leakage is sufficient to provide a pressure drop across the diaphragm 60 in a downstream direction, which during cruise of an associated aircraft (not shown) urges the diaphragm 60 against the protuberance 58 and thus prevents communication between the spaces 50 and 52.

Referring now to Fig. 3. In operation of the gas turbine engine (not shown) during take off of an associated aircraft (not shown), the temperature experienced by the turbine 10 is higher than at cruise of the aircraft (not shown). Temporarily therefore, more cooling air is needed over and above that used at cruise, and is provided as follows.

During take off of the aircraft (not shown) a thermocouple 66 senses the turbine temperature and passes the resulting voltage signal to a black box 68. The signal is processed so as to enable it to actuate the solenoid valve 64 and thus vent the pressurised air from the space 51 on the upstream side of the diaphragm 60. Air passing into the space 50 thus flexes the diaphragm away from the annular land 58 and consequently flows past the land 58, through space 52, the passageway 36 and the swirlers 42, into the plenum space 38, and thence through apertures 39 in the disc 28 to the roots of the turbine blades 20.

The air thereafter enters the turbine blades 20 through apertures (not shown) and cools them in known manner.

Referring now Fig. 4. The embodiment depicted thereby consists of local bosses 72, only one of which is shown, which are equiangularly spaced around the hub of the member 32 (Fig. 1). Each boss 72 is internally relieved to provide a space 74 which contains an annular rim 76 against which a disc like flexible diaphragm 77 seats. A cap 78 is provided for each boss 72 and when fitted locates the rim of the diaphragm 77 between the internal rim 76 and the rim 80 of the cap 78. The cap 78 is held in position by peening the rim 82 of the boss 72 over the edge of the sloping shoulder 84-of the cap 78.

The cap 78 includes two internal spaces 86 and 88 which are seperated by a diametrically arranged rib 89. A circular rim however, could be substituted therefore. An air inlet 92 and an air outlet 94 are provided to space 86 and from space 88 respectively. As in the one operating mode described hereinbefore with respect to Figs. 1 and 2, the diaphragm 77 in the embodiment of Fig. 4 seals against the rib 89 to prevent communication between the inlet 92 and the outlet 94. Pressure on the upstream face of the diaphragm 77 so to maintain the seal, is, as before, provided by the pressure drop across the diaphragm 77.

In both embodiments described herein, when the associated aircraft (not shown) reaches its cruise altitude, the engine (now shown) is throttled back, the turbine temperature falls and the signal is removed from the valve 64 (Fig. 3) The diaphragm 60 or 77 is urged against the annular land 58 or the rib 90 and so cuts off the cooling air supply through the air inlets and air outlets 50 and 52, or 92 and 94.

Each air outlet 94 is connected via ducting not shown in Fig. 4, but which corresponds to the structure 36, 42 and 39 of Fig. 1, to the roots of turbine blades 20.

An alternative embodiment (not shown) consists of connecting the pilots' throttle lever to the valve 64. The connection could be direct, or electrical in which case the valve 64 could again be solenoid operated. In both cases, the valve 64 will be moved to vent the space 51, when the throttle is opened to achieve take off.

Claims

1. A gas turbine engine including a valve device for controlling a flow of cooling air to a turbine stage, comprising a chamber, one internal wall surface of which has a protuberance, a diaphragm which in one position seals against the chamber rim and the protuberance, and a further space on the other side of the diaphragm, cooling air inlet means on one side of the protuberance, cooling air outlet means on the other side of said protuberance and connected to a turbine cooling flow path and means for enabling flexing of the diaphragm away from the protuberance and so enable communication between said air inlet means and said air outlet means, said flexing means being adapted for actuation at the onset of an operating regime which results in an increase in operating temperature of the turbine stage of the gas turbine engine.

2. A gas turbine engine including a valve as claimed in claim 1 wherein the means for enabling flexing of the diaphragm comprises a removable counter pressure for, prior to actuation, countering air pressure on the air inlet side of the diaphragm so as to in one mode of operation, maintain the diaphragm seal

3. A gas turbine engine including a valve as claimed in claim 2 wherein the counter pressure comprises a pressurized air supply in communication with said further chamber and includes means for venting said air supply, on the turbine stage achieving said operations regime.

4. A gas turbine engine including a valve as claimed in claim 3 wherein the means for venting said air supply comprises a bleed point in the air supply and a further valve blocking the bleed point and movable to unblock said bleed point in response to a signal generated on achievement of said increase in temperature of said turbine stage.

5. A gas turbine engine including a valve as claimed in claim 4 wherein the further valve comprises a solenoid which is connected for actuation in response to said signal.

6. A gas turbine engine including a valve as claimed in claim 4 wherein the said further valve is connected for actuation by the throttle of the gas turbine engine.

7. A gas turbine engine including a valve as claimed in any previous claim, wherein said chamber is annular and includes a plurality of said air inlets and said air outlets which are equi-angularly around the annulus.

8. A gas turbine engine including a valve as claimed in any of claims 1 to 6 comprising a plurality of said chambers equi-angularly spaced in and around the downstream face of an annular member.

9. A gas turbine engine including a valve device as claimed in any previous claim wherein each air inlet has a cooperating air outlet positioned radially inwardly thereof.

10. A gas turbine engine including a valve device as claimed in claim 8 wherein respective, cooperating air inlets and air outlets are separated by an annular protuberance.

11. A gas turbine engine including a valve device substantially as described in this specification and with reference to Figs. 1 to 3 of the drawings.

12. A gas turbine engine substantially as described in this specification and with reference to Fig. 4 of the drawings.