GB2555379A

GB2555379A - Gas turbine engine heat exchanger

Info

Publication number: GB2555379A
Application number: GB1617638.0A
Authority: GB
Inventors: Bond Jonathan; Bewick Claire; Hussain Zahid
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2016-10-18
Filing date: 2016-10-18
Publication date: 2018-05-02
Also published as: GB201617638D0

Abstract

A ducted fan gas turbine engine 10 having a bypass duct 22 surrounding a wall structure 26 that encases a core accessory zone 27 containing core engine accessories (e.g. accessory gearbox, turbine case cooling system, bleed valves, combustor manifold, variable stator vanes actuators, etc.). The core accessory zone 27 surrounds a core engine and fluidly communicates with an external vent 28. Air pressure in the core accessory zone is at or above ambient pressure. A heat exchanger (31, figure 2, e.g. matrix OAHE) is located within a cooling duct 25 for cooling an engine fluid (e.g. lubricating oil or generator oil) by heat exchange with air flow, diverted from bypass duct 22 through an air inlet 29 formed in the wall structure 26. The diverted flow is delivered to the core accessory zone 27 downstream of the heat exchanger 31 before exiting through vent 28. One or more aerosol capture elements (33) may comprise gauze barriers (39a, 39b, figure 4), an inertial separator having staggered projecting members (35a, 35b, figure 2) and oil drainage hole (37) and a bend of more than 90 degrees for reducing fire risk by isolating heat exchanger 31 oil leakage from the accessory zone 27.

Description

(54) Title of the Invention: Gas turbine engine heat exchanger

Abstract Title: Gas turbine engine heat exchanger located within a cooling duct (57) A ducted fan gas turbine engine 10 having a bypass duct 22 surrounding a wall structure 26 that encases a core accessory zone 27 containing core engine accessories (e.g. accessory gearbox, turbine case cooling system, bleed valves, combustor manifold, variable stator vanes actuators, etc.). The core accessory zone 27 surrounds a core engine and fluidly communicates with an external vent 28. Air pressure in the core accessory zone is at or above ambient pressure. A heat exchanger (31, figure 2, e.g. matrix OAHE) is located within a cooling duct 25 for cooling an engine fluid (e.g. lubricating oil or generator oil) by heat exchange with air flow, diverted from bypass duct 22 through an air inlet 29 formed in the wall structure 26. The diverted flow is delivered to the core accessory zone 27 downstream of the heat exchanger 31 before exiting through vent 28. One or more aerosol capture elements (33) may comprise gauze barriers (39a, 39b, figure 4), an inertial separator having staggered projecting members (35a, 35b, figure 2) and oil drainage hole (37) and a bend of more than 90 degrees for reducing fire risk by isolating heat exchanger 31 oil leakage from the accessory zone 27.

Fig. 1

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Fig. 2

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Fig. 3

Fig. 4

GAS TURBINE ENGINE HEAT EXCHANGER

Field of the Invention

The present invention relates to heat exchange arrangements within gas turbine engines.

Background

Heat management is an important consideration for the design of a gas turbine engine. Efficient arrangements for heat generation and rejection within the engine maintain oil system and fuel system temperatures within their respective temperature operational limits, while ensuring low engine performance loss.

For example, the oil used to lubricate the bearing chambers and gearboxes on a turbofan engine needs to be maintained below a thermal limit. Typically, this is accomplished using the engine fuel as a heat sink, but on some engines this is insufficient and heat needs be lost to another sink, for example air. Conventionally this can be achieved by, for example, surface heat exchangers fitted on the rear fan case or surface heat exchangers fitted on the bypass duct inner surface. However, such heat exchangers introduce flow-disturbing features into airflows in the bypass duct and may also reduce the surfaces available for installing acoustic liners.

Matrix heat exchangers fed by offtakes from the bypass air flow have also been used, but as bypass ratios increase and fan pressure ratios decrease, the pressure gradients available for such conventional matrix systems is reduced.

In particular, newer high bypass ratio engines tend to have small cores and shorter rear fan cases, which reduce the available space for heat management hardware. For example: the area for surface heat exchangers is reduced and the relative cost of the pressure loss they induce in the bypass duct is increased; a short rear fan case makes a bypass duct to fan cowl door exit matrix heat exchanger impractical to install; and pressure gradients to drive flow through core engine mounted matrix heat exchangers are reduced.

Moreover, engines incorporating a power gearbox require even more oil cooling capacity than non-geared turbofans.

Accordingly, there is a need to provide improved cooling systems that can be packaged in a compact manner within the gas turbine engine.

Summary

The present inventors have realised that there is the potential to have efficient heat exchanger that can be packaged in a compact manner by exhausting the coolant air into the core accessory zone of the engine. The core accessory zone on e.g. a large civil turbofan is generally at a pressure close to ambient, and hence a significant pressure gradient is available between an offtake from the bypass duct and the core accessory zone to drive airflow through a heat exchanger. This arrangement is also advantageous from a noise perspective as any noise created by the coolant air exhaust can be shielded by the fairing encasing the core accessory zone. In addition, the number of cut-outs in any bypass duct noise liner can be reduced.

Thus the present invention provides a ducted fan gas turbine engine having a core engine, a core accessory zone surrounding the core engine and containing core engine accessories, a wall structure encasing the core accessory zone and the core engine, and an annular bypass duct surrounding the wall structure, wherein:

the core accessory zone fluidly communicates with an external vent such that the air pressure in the core accessory zone is at or above ambient pressure;

a cooling duct having an air inlet formed in the wall structure is provided such that, during operation of the gas turbine engine, an air flow is diverted through the air inlet and into the cooling duct from a bypass air flow through the bypass duct;

a heat exchanger is located within the cooling duct for cooling of an engine fluid by heat exchange with the diverted air flow; and downstream of the heat exchanger the cooling duct delivers the diverted flow to the core accessory zone such that the diverted air flow mixes with air used to bathe the core engine accessories before leaving the core accessory zone through the vent.

The present invention therefore enables a compact and efficient cooling system, in which the pressure difference between the inlet of the cooling duct and the core accessory zone can drive air flow through the heat exchanger. As the inlet can be of relatively small size, acoustic liner loss in the bypass duct can be reduced.

Furthermore, by exploiting the pressure difference between the cooling duct inlet and the core accessory zone, the need for ejectors, or similar flow-enhancing devices, may be avoided.

Optional features of the invention will now be set out. These are applicable singly or in any combination.

The bypass duct may contain outlet guide vanes, and the inlet to the cooling duct may be formed downstream of these vanes. The inlet may be dedicated to supply the cooling duct and heat exchanger with the diverted flow of bypass air, or alternatively the inlet may feed e.g. a plenum from which heat exchangers and other systems, for example turbine case cooling and ventilation, are also fed. In such arrangements, it will be understood that only a portion, rather than all of, of the diverted flow may be used for cooling.

Conveniently, the vent from the core accessory zone may be located in the wall structure, e.g. downstream of the bypass exhaust nozzle which terminates the annular bypass duct. Another option, however, is for the vent to be located in an outer surface of a nacelle which surrounds the bypass duct, air from the core accessory zone being suitably ducted to the vent, e.g. via piping extending through outlet guide vanes contained in the bypass duct. The vent is typically sized such that the air pressure in the core accessory zone is similar to the pressure at the vent, which is typically slightly above ambient pressure. In this way a suitable pressure gradient from the cooling duct inlet to the core accessory zone can be provided.

The core engine accessories may include any one or any combination of: an accessory gearbox and associated unit(s), turbine case cooling system(s), bleed valve(s), combustor manifold(s), secondary air system(s), pneumatic solenoid(s) and/or variable stator vanes actuator(s).

The heat exchanger may be a matrix heat exchanger, e.g. of tube or plate construction. The diverted air flow which passes through the heat exchanger is heated by the heat exchange with the hotter engine fluid and, as such, the air flow exhausting from the heat exchanger may be suitable for use in subsequent heating of other parts of the gas turbine engine. For example, the exhausted air flow can be used for icing prevention of e.g. pressure sense pipes.

The heat exchanger may be adapted to receive core engine lubricating oil such that the lubricating oil is cooled by heat exchange with the diverted air flow. However, the heat exchanger may alternatively be adapted to receive other engine fluids for cooling, such as e.g. generator oil.

To reduce any fire risk from accidental egress of engine fluid to the core accessory zone due to e.g. leakage of the fluid from the heat exchanger, it may be desirable to provide further, optional features to assist in reducing or mitigating fire risk. Such features may be particularly desirable in the case of engine oil leaking in the form of an inflammable of oil aerosol. In particular, one or more aerosol capture elements may be located downstream of the heat exchanger in the cooling duct, the aerosol capture elements leakage isolating the heat exchanger from the core accessory zone. Aerosols formed by leakage from the heat exchanger will tend to flow in the air flow direction through the duct, and so positioning the capture elements downstream of the heat exchanger promotes their capture. Leakage isolation may include preventing all, substantially all, or the majority of any aerosol droplets produced by leakage from the heat exchanger from entering the core accessory zone.

The aerosol capture elements may include a gauze barrier extending across the cooling duct. The mesh of the gauze may be selected as appropriate to allow a suitable air flow through the duct, whilst providing sufficient leakage isolation of the heat exchanger from the core accessory zone. It will be appreciated that there is a trade-off between these two factors. Use of a gauze barrier offers the additional advantage that it may provide improved fire containment in the case of a fire within the duct. This is due to the “Davy lamp” principle, the gauze acting as a flame arrestor by virtue of a mesh size which is too small to allow propagation of a flame past the gauze. Indeed, a further gauze barrier may be located upstream of the heat exchanger in the cooling duct, and extending across the cooling duct. The upstream barrier can further assist in capturing any aerosols which are released in an upstream direction from a leaking heat exchanger, and may provide an additional flame arrestor on the other side of the heat exchanger.

The aerosol capture elements may include an inertial separator. For example, the inertial separator may include two or more staggered projecting members which project into the cooling duct to prevent line-of-sight flow of aerosols through the duct from the heat exchanger to the core accessory zone. The size and shape of these projecting members are not particularly limited.

The inertial separator may additionally or alternatively include a bend in the cooling duct, e.g. of 90° or more. In some cases, the bend may be 180° or more. In this way, the inertia of the aerosol droplets will tend to cause the droplets to impinge on the wall of the cooling duct at the bend, thus removing the aerosol droplets from the air flow through the duct, and thus leakage isolating the heat exchanger from the core accessory zone.

The part(s) of the inertial separator on which the aerosol droplet impinge (i.e. duct walls, projecting members) may be termed “wet-out surfaces”. Upon impingement with these wetout surfaces, the aerosol droplets will tend to coalesce, and may be removed from the duct by provision of e.g. drainage channels or pipes in the duct into which the coalesced fluid may collect. In this way, pooling of inflammable fluid in the duct can be avoided. The removed fluid may by led to a tell-tale drains mast to detect oil leaks. Alternatively or additionally, a lipped drip tray may be provided to catch the coalesced fluid. One or more drainage holes may be provided in the duct wall to assist in removal of the fluid.

Brief Description of the Drawings

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:

Figure 1 shows a longitudinal cross-section through a ducted fan gas turbine engine;

Figure 2 shows a schematic longitudinal cross-section through a front section of a wall structure of the engine of Figure 1;

Figure 3 shows a schematic longitudinal cross-section through a cooling duct having a variant aerosol capture element; and

Figure 4 shows a schematic longitudinal cross-section through a cooling duct having a further variant aerosol capture element.

Detailed Description and Further Optional Features

With reference to Figure 1, a ducted fan gas turbine engine incorporating the invention is generally indicated at 10 and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23. However, the present invention is not limited to the engine architecture shown in Figure 1. For example, the present invention may also be applied to two-shaft or geared engines.

During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate-pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediatepressure compressor 13 compresses the air flow A directed into it before delivering that air to the high-pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the exhaust nozzle 19 to provide additional propulsive thrust. The high, intermediate and lowpressure turbines respectively drive the high and intermediate-pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

A wall structure 26 encases a core accessory zone 27 surrounding the core engine and containing core engine accessories. The annular bypass duct 22 surrounds the wall structure. The wall structure has a vent 28 to the rear of the core accessory zone which equalises the air pressure in the core accessory zone to the air pressure outside the wall structure at the vent. This brings the pressure in the core accessory zone to slightly above ambient. A cooling duct 25 having an air inlet 29 formed in the wall structure is provided such that, during operation of the gas turbine engine, an air flow is diverted into the cooling duct from the second air flow B flowing through the bypass duct 22. The diverted air flow is then exhausted from the cooling duct into the front of the core accessory zone, whereupon it courses around the engine accessories in the zone before exiting at the vent 28. The pressure difference between the air pressure at the air inlet to the cooling duct and the air pressure in the core accessory zone drives the diverted air flow through the duct.

Rather than exhausting the diverted air flow to the front of the core accessory zone 27, the cooling duct 25 can be extended to exhaust the diverted air flow to a more rearward position in the core accessory zone, i.e. closer to the vent 28. However, if the duct is extended in this way, before leaving the wall structure through the vent the diverted air flow still mixes with the air used to bathe the core engine accessories, such that the pressure difference driving the diverted air flow is maintained.

Figure 2 shows a schematic longitudinal cross-section through a front section of the wall structure 26, showing an approximate location of the cooling duct 25. The inlet 29 to the duct is downstream of outlet guide vanes 30 of the engine 10. The diverted air flow into the duct is indicated with an arrow. A matrix heat exchanger 31 is located within the cooling duct, and is adapted to receive a flow of core engine lubricating oil so that, during operation, the core engine lubricating oil is cooled by heat exchange with the relatively cool diverted air flow. The cooling duct and heat exchanger provide a compact and efficient cooling system which does not significantly reduce the surfaces available for installing acoustic liners in the bypass duct 22 and does not introduce significant flow-disturbing features into the bypass duct. Moreover, it is also well shielded by the wall structure, and thus does not contribute substantially to engine noise.

An optional aerosol capture element in the form of an inertial separator 33 is located in the cooling duct 25 downstream of the heat exchanger, to reduce any fire risk in the event of leakage of oil from the heat exchanger.

The inertial separator 33 forces the air flow along a convoluted path and includes two staggered projecting members 35a, 35b projecting into the cooling duct 25 to provide wetout surfaces on which aerosol droplets in the air flow may impinge. In particular, the projecting members prevent line-of-sight flow of aerosols through the cooling duct from the heat exchanger to the core accessory zone, thus providing leakage isolation of the heat exchanger and the core accessory zone. A drainage hole 37 can be provided in the wall of the cooling duct. This can assist in removal of liquid resulting from coalescence of captured aerosol droplets from the cooling duct. Whilst two projecting members are shown in this embodiment, the number of projecting members is not particularly limited. More projecting members may improve the leakage isolation effectiveness of the aerosol capture element.

Figure 3 shows a schematic longitudinal cross-section through a cooling duct 25 having a second variant aerosol capture element. The aerosol capture element is an inertial separator 33 comprising a deviation in duct direction of 180° downstream of the matrix heat exchanger 31. This bend in the duct acts as a trap so that part of the duct wall forms a wetout surface on which aerosol droplets in the airflow may impinge. It is not necessary that the deviation in duct direction is exactly 180°. The deviation in duct direction may be larger or smaller than 180°. However, the deviation in duct direction is preferably at least 90° in order to provide sufficient leakage isolation of the heat exchanger from the core accessory zone.

Such inertial separators 33 can be located at positions where the cooling duct 25, in operation, is at a temperature lower than the auto-ignition temperature of oil. Thus any captured and drained oil can be safely collected and disposed of. However, Figure 4 shows a schematic longitudinal cross-section through a cooling duct 25 having a further variant aerosol capture element. Here the aerosol capture element includes a gauze barrier 39a extending across the cooling duct downstream of the heat exchanger 31. Another gauze barrier 39b is also located upstream of the heat exchanger in the cooling duct and extending across the cooling duct. The size of the mesh of the gauze may be selected to allow a suitable air flow through the duct, whilst providing a sufficient barrier to passage of aerosol droplets from the heat exchanger. The gauze barriers can also act as flame arrestors to provide fire containment in the case of a fire within the duct.

While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

Claims

1. A ducted fan gas turbine engine (10) having a core engine, a core accessory zone (27) surrounding the core engine and containing core engine accessories, a wall structure (26) encasing the core accessory zone and the core engine, and an annular bypass duct (22) surrounding the wall structure, wherein:

the core accessory zone fluidly communicates with an external vent (28) such that the air pressure in the core accessory zone is at or above ambient pressure;

a cooling duct (25) having an air inlet (29) formed in the wall structure is provided such that, during operation of the gas turbine engine, an air flow is diverted through the air inlet and into the cooling duct from a bypass air flow through the bypass duct;

a heat exchanger (31) is located within the cooling duct for cooling of an engine fluid by heat exchange with the diverted air flow; and downstream of the heat exchanger the cooling duct delivers the diverted flow to the core accessory zone such that the diverted air flow mixes with air used to bathe the core engine accessories before leaving the core accessory zone through the vent.

2. The gas turbine engine of claim 1, wherein the heat exchanger is adapted to receive core engine lubricating oil, which is cooled by heat exchange with the diverted air flow.

3. The gas turbine engine of claim 1 or 2, wherein the vent is located in the wall structure downstream of a bypass exhaust nozzle (23) which terminates the annular bypass duct.

4. The gas turbine engine of any one of the previous claims, wherein one or more aerosol capture elements (33) are located in the cooling duct downstream of the heat exchanger, the aerosol capture elements leakage isolating the heat exchanger from the core accessory zone.

5. The gas turbine engine of claim 4, wherein the one or more aerosol capture elements include a gauze barrier (39a) extending across the cooling duct.

6. The gas turbine engine of claim 5, wherein a further gauze barrier (39b) is located upstream of the heat exchanger in the cooling duct and extends across the cooling duct.

7. The gas turbine engine of any one of claims 4 to 6, wherein the one or more aerosol capture elements include an inertial separator.

8. The gas turbine engine of claim 7, wherein the inertial separator includes two or more staggered projecting members (35a, 35b) which project into the cooling duct to prevent lineof-sight flow of aerosols through the duct from the heat exchanger.

9. The gas turbine engine of claim 7 or 8, wherein the inertial separator includes a bend 5 in the cooling duct of 90°.

10. The gas turbine engine of any one of the previous claims, wherein the heat exchanger is a matrix heat exchanger.

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Application No: Claims searched:

GB1617638.0

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