GB2517647A

GB2517647A - Gas turbine engine

Info

Publication number: GB2517647A
Application number: GB8035432.7A
Authority: GB
Inventors: Geoffrey Stephen Hough
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1980-11-04
Filing date: 1980-11-04
Publication date: 2015-03-04
Anticipated expiration: 2000-11-04
Also published as: GB2517647B; GB8035432D0

Abstract

A gas turbine engine comprises an annular array of nozzle guide vanes 24, each guide vane including a heat pipe which extends into an annular cooling air duct 19 located radially inwardly of the engine's combustion chamber 12. After passing over the heat pipes and thereby cooling the nozzle guide vanes, the cooling air passing through the duct is directed on to the radially inner part 36 of the aerofoil portion 26a of the first stage of rotor blades 26 downstream of the nozzle guide vanes. The cooling air is directed on to the rotor blades at such an angle that the blades extract useful work from the cooling air.

Description

ROLLS-ROYCE LTh CASE NO.

SHORt TITLE: "KEAT PIPE NGV WITH HEAT EXCHANGER" APPLICATION NO.

DATED: 1 APPLICATION - -PATENt'S AOl' 1977

SPECIFICATION

GAS TURBINE ENGINE

This invention relates to gas turbine engines and in particular to the combustion equipment and turbines of such engines.

The hot gases which are issued in operation from the combustion chamber or chambers of a gas turbine engine are conventionally directed onto the first annular array of rotor aerofoil blades of the turbine of the engine by an annular array of vanes which are usually referred to as nozzle guide vanes. The gases from the combustion chamber are, at this point, extremely hot. Consequently special precautions have to be taken in order to ensure that the temperatures of the nozzle guide vanes are not allowed to exceed their operational limits.

This is frequently achieved by passing a cooling fluid, usually air, through passages within the vanes. Such cooling does however impose penalties upon the overall efficiency of the engine by cooling down the efflux gases from the combustion chamber and necessitating the use of nozzle guide vane structures having complicated internal cooling passageways which are difficult and expensive to produce.

It is an object of the present invention to provide a gas turbine engine in which nozzle guide vane cooling is achieved by means which impose reduced penalties upon overall engine efficiency and which does not require the provision of complicated cooling passages within the nozzle guide vanes.

According to the present invention, a gas turbine engine includes a compressor, a turbine having at least one annular array of rotor blades each having an aerofoil cross-section portion, combustion equipment, an annular array of nozzle guide vanes each having an aerofoil cross-section portion and so disposed as to direct the hot gas efflux issued in operation from said combustion equipment onto said annular array of rotor blades, and duct means interconnecting said compressor and said turbine, said duct means being adapted in operation for the passage therethrough of cooling air derived from said compressor and so disposed as to direct that cooling air onto a part of the aerofoil cross-section portion of each of said rotor blades, each of said nozzle guide vanes comprising at least one heat pipe so disposed as to constitute at least a part of the aerofoil portion of its respective nozzle guide vane and adapted to extend into said duct means in such a manner as to be in heat exchange relationship with said cooling air passing in operation therethrough.

Said duct means is preferably so disposed as to direct said cooling air onto a part of the aerofoil cross-section portion of each of said rotor blades at such an angle that said rotor blades extract useful work from said cooling air.

Said duct means may be so disposed as to direct said cooling air onto the radially inner part of the aerofoil cross-section portion of each of said rotor blades.

Said duct means may be partially defined by the radially inner portion of said combustion equipment.

The whole of the aerofoil cross-section portion of each of said nozzle guide vanes may be constituted by a single heat pipe.

Those portions of said heat pipes adapted to extend into said duct means may be of generally aerofoil-shaped cross-section.

Splitter means may be provided in said duct means between adjacent aerofoil shaped cross-section heat pipe portions and are so configured as to direct the majority Qf cooling air flowing in operation through said duct means into close proximity with said heat pipe portions.

Said heat pipe portions contained within said duct means preferably complitely extend across said duct means.

Said duct means and said combustion equipment are preferably annular.

The invent Ion will now be particularly described by way of example with reference to the accompanying drawings in which:-Figure 1 is a side view of a gas turbine engine in accordance with the present invention.

Figure 2 is a sectional side view of a portion of the combustion equipment and turbine of the gaa turbine engine shown in Figure 1.

Figure 3 is a view on section line A-A of Figure 2.

Figure 4 is a view on section line B-B of FIgure 2.

With reference to Figure 1 an axial flow gas turbine engine generally indicated at 10 comprises, in flow series, a compressor section 11, combustion equipment 12, a turbine section 13 and a propulsion nozzle 14. The gas turbine engine 10 functions in the conventional manner, that is, air compressed in the compressor section 11 is mixed with fuel and the mixture combusted in the combustion equipment 12 and expanded through the turbine 13 to atmosphere via, the propulsion nozzle 14.

The combustion equipment 12, as can be seen in more detail in Figure 2, comprises an annular combustion chamber 15 having a number of inlets 16 which are adapted to receive fuel vapourisers of the conventional type (not shown in the interest ? simplicity). The combustion chamber 15 is enclosed by radially inner and outer casing structures 17 and 18 respectively so that radially inner and outer annular ducts 19 and 20 respectively are defined by both the casing structures 17 and 18 and the combustion chamber 15. The casing structures 17 and 18 converge upstream of the combustion chamber inlets 16 to provide a further inlet 20 which is so positioned as to receive compressed air from the compressor 11 (not shown in Figure 2). The air from the compressor 11 flows in the direction generally indicated by the arrow 21.

After entering the inlet 20 the air flow, is divided with one portion flowing through the inlet 16 and into the combustion chamber 15 and the remainder flowing into the annular ducts 19 and 20. The air which enters the combustion chamber 15 is mixed with fuel and the mixture combusted in the conventional manner. General cooling of the combustion chamber 15 is achieved by the passage of cooling air throu&a the annular ducts 19 and 20. More specific cooling of those areas of the combustion chamber 15 which reach the highest temperatures is achieved by the provision of numerous small ducts 22 which interconnect the interior of the combustion chamber 15 with the annular ducts 19 and 20. These small ducts ensure en adequate cooling air supply to those areas which require cooling. In addition to the small ducts 22, the wall of the combustion chamber 15 is provided with somewhat larger ducts 23 which provide air flows to those areas of the combustion chamber 15 where additional air is required for satisfactory combustion.

At the downstream end of the combustion chamber 15, there is provided an annular array of nozzle guide vanes 24. Each of the nozzle guide vanes 24 is provided with an aerofoil cross-section portion 25 which is so disposed that together the nozzle guide varies 24 direct the efflux gases from the combustion chamber 15 onto an annular array of rotor blades 26 each having an aerofoil cross-section portion 26a. The rotor blades 26 are mounted on a disc 27 which *is rotatable about the engine axis lOa and comprise a portion of the turbine 13 of the engine 10.

Each ofthe nozzle guide vanes 24 is provided with radially inner and outer shrouds 28 and 29 respectively which cooperate with the ceresponding shrouds 29 and 28 of adjacent nozzle guide vanes 24 to define downstream extensions of the walls of the combustion chamber 15.

The aerofoil cross-section portion 25 of each nozzle guide vane 24 extends inwardly of its respective radially inner shroud 28 to terminate in a further shroud 30. The further shrouds 30 of adjacent nozzle guide vanes 24 cooperate to define a downstream extension of the radially inner casing structure 17 so that the shrouds 28 and define a downstream extension of the annular duct 19.

That extension 31 of each nozzle guide vane 24 which extends between the inner shroud 28 and the further shroud 30, and the aerofoil shaped portion 25 of each nozzle guide vane 24 are hollow as can be seen in Figures 3 and 4. They define a common chamber 32 which is sealed and evacuated, and provided on its wall with a stainless steel mesh wick structure 33. The chamber 32 additionally contains a small amount of sodium so that the aerofoil cross-section portion and the nozzle guide vane extension 31 together constitute a single heat pipe.

Throughout this specification, the term "heat pipe" is to be understood as meaning a heat transfer device cothprising a sealed container which encloses both a condensable vapour and capillary means capable of causing the transport of the condensed vapour from a cooler area of the container to a hotter area, the condensable vapour being transported from the hotter area to the cooler area by the vapour pressure gradient between the two areas, the vapour being condensed in the cooler area.

Thus since the aerofoil cross-section portion 25 and the nozzle guide vane extension 31 constitute a single heat pipe, then they have an extremely high thermal conductivity so that under steady state operating conditions, they are effectively isothermal.

This means that in operation, the hot gases which constitute the ef flux from the combustion chamber 15 pass over the aerofoil cross-section portions 25 of the nozzle guide 24 and heat up those portions which then in turn heat up the nozzle guide extensions 31.

There is, as previously stated, a flow of cooling air through the annular duct 19 which serves to provide combustion chamber cooling... However since the nozzle guide vane extensions 31 extend across this duct 19, then they too are subject to cooling by the air flow. As each nozzle guide vane extension 31 and its corresponding aerofoil portion 25 constitute portions of a single heat pipe then the cooling of extensions 31 results in turn in the cooling of the aerofoil cross-section portions 25.

In order to ensure adequate cooling of the nozzle guide vane extensions 31, their surfaces are provided with a large number of protuberances 34 so as to increase their effective surface areas.

Noreover, splitte' members 35 are positioned between adjacent nozzle guide vane extensions 31 so as to ensure that as much as possible of the cooling air passing in operation through the annular duct 19 is brought into heat exchange relationship with the nozzle guide vane extensions 31.

The inner and further shrouds 28 and 30 respectively of the nozzle guide vanes 24 are so configured as to define together an annulus of decreasing cross-sectional area (in a downstream direction) which directs the cooling air from the annular duct 19 onto the radially inner parts 36 of the aerofoil cross-section portions 26a of the rotor blades 26. Ihus the downstream ends of the further shrouds 30 are aligned with the radially inner shrouds 37 of the rotor blades 26 whilst the downstream ends of the nozzle guide vane inner shrouds 28 teninate somewhat radially outward of this position. Consequently air which has served to cool the nozzle guide vanes 24 and part of the combustion chamber 15 serves the additional function of providing cooling of the radially inner parts 36 of the aerofoil cross-section portion 26a of the rotor blades 26, an area which is particularly prone to high stress levels as a result of reaching high temperatures. The cooling air stream from the annular duct 19 also provides a barrier between the hot gas stream from the combustion chamber 15 and the upstream face of the disc 27. Consequently if any gas leakage occurs down the upstream face of the disc 27, it is of the relatively cool air stream from the annular duct 19 and not the hot gas stream.

Since such leakage is not as serious as a hot gas stream leakage, the sealing system at the junction between the radially outer region of the disc 27 and the downstream radially inner region of each nozzle guide vane 24 need not be as complicated as is the case in conventional arrangements.

Phe splitter members 35, the nozzle guide vane extensions 31 and -the shrouds 28 and 30 are so configured as to direct the cooling air stream from the duct 19 onto the aerofoil cross-section portions 26a of the rotor blades 26 so that the rotor blades 26 extract useful work from the cooling air stream. In a typical case in which the velocity of the cooling air stream is about 75% of that of the hot gas stream, the angle at which the cooling air stream is directed onto the rotor blades 26 is greater than that at which the hot gas stream is so directed by the nozzle guide vane aerofoil portions 25.

It may be that in certain circumstances, it is the radially outer regions of the rotor blades 26 which are particularly prone to thermal stress problems. In such a situation, the nozzle guide vanes 24 could be so configured that their extensions 31 project into the annular duct 20 radially outwardly of the combustion chamber 15.

Suitably shaped shrouds could then be provided to direct the cooling air stream from the duct 20 on the radially outer regions of the rotor blades 26.

The present invention therefore provides nozzle guide vane cooling without the need for complicated passageways within the vanes.

Moreover the air used for cooling the nozzle guide vanes serves the additional functions of a limited degree of rotor blade cooling as well as perfoSing work on the rotor blades. -8.-

Claims

What we claim is:- 1. A gas turbine engine including a compressor, a turbine having at least one annular array of rotor blades each having an aerofoil cross-section portion, combustion equipment, an annular array of nozzle guide vanes each having an aerofoil cross-section portion and so disposed as to direct the hot gas efflux issued in operation from said combustion equipment onto said annular array of rotor blades and duct means interconnecting said compressor and said turbine, said duct means being adapted in operation for the passage therethrough of cooling air derived from said compressor and so disposed as to direct that cooling air onto a part of the aerofoil cross-section portion of each of said rotor blades, each of said nozzle guide vanes comprising at least one heat pipe so disposed as to constitute at least a part of the aerofoil cross-section portion of its respective nozzle guide tane and adaited to extend into said duct meaxis in such a manner as to be in heat exchange relationship with cooling air passing in operation therethroui.
2. A gas turbine engine as claimed in claim 1 wherein said duct is so disposed as to direct said cooling air onto a part of the aerofoil cross-section portion of each of said rotor blades at such an angle that said rotor blades extract useful work from said cooling air.
3. A gas turbine engine as claimed in claim 1 or claim 2 wherein said duct means is so disposed as to direct said cooling air onto the radially inner part of the aerofoil cross-section portion of each of said rotor blades.
4. A gas turbine engine as claimed in any one preceding claim wherein said duct means is partially defined by the radially inner portion of said combustion equipment.
5. A gas turbine engine an claimed in any one preceding claim wherein the whole of the aerofoil cross-section portion of each of said nozzle guide vanes is constituted by a single heat pipe.
6. A gas turbine engine a claimed in any one preceding claim wherein those portions of said heat pipe adapted to extend into said duct means are of generally aerofoil-shaped cross-section. -.9-
7. A gas turbine engine as claimed in claim 6 wherein splitter means are provided in said duct means between adjacent aerofoil shaped cross-section heat pipe portions and are so configured as to direct the majority of cooling air flowing in operation through said duct means into close proximity with said heat pipe portions.
8. A gas turbine engine as claimed in any one preceding claim wherein said heat pipe portions contained within said duct means completely extend across said duct means.
9. A gas turbine engine as claimed in any one preceding claim wherein said duct means and said combustion equipmeflt are annular.
10. A gas turbine engine substantial1,r as hereinbefore described with reference to and as shown in the accompanying drawings.