Title:
INSTALLATION HAVING A THERMAL TRANSFER ARRANGEMENT
Kind Code:
A1


Abstract:
A gas turbine engine 2 is installed in a nacelle 4 to define a bay 16 which is ventilated by air entering the bay 16 through an intake 27. A partition 28 in the bay 16 includes a permeable wall 34 provided with flow passages 36 which direct the air flow in the bay 16 towards a turbine casing 14. The air flow is emitted from the flow passages 36 as jets 46 to provide impingement cooling of the turbine casing 14.



Inventors:
Avenell, Christopher S. (Bristol, GB)
Application Number:
13/081938
Publication Date:
10/27/2011
Filing Date:
04/07/2011
Assignee:
ROLLS-ROYCE PLC (London, GB)
Primary Class:
Other Classes:
165/47
International Classes:
F28D20/00; F01D25/12
View Patent Images:
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Primary Examiner:
MCCAFFREY, KAYLA M
Attorney, Agent or Firm:
OLIFF PLC (ALEXANDRIA, VA, US)
Claims:
1. An installation comprising a component having a means for inducing fluid flow through the component, and a thermal transfer arrangement, the thermal transfer arrangement comprising a chamber which is ventilated in operation of the installation by a forced fluid flow induced by a means different to the means for inducing fluid flow through the component, the chamber being bounded at least partly by a heat transfer surface of the component, a flow passage being disposed within the chamber, which flow passage is oriented to direct at least part of the fluid flow at the heat transfer surface to transfer heat between the fluid and the heat transfer surface.

2. An installation comprising a component having a means for inducing fluid flow through the component, and a cooling arrangement for cooling the component, the cooling arrangement comprising a chamber which is ventilated in operation of the installation by a forced fluid flow induced by a means different to the means for inducing fluid flow through the component, the chamber being bounded at least partly by a surface of the component, a flow passage being disposed within the chamber, which flow passage is oriented to direct at least part of the fluid flow at the surface of the component, to cool the component.

3. An installation as claimed in claim 2, in which the flow passage is provided in a wall situated within the chamber.

4. An installation as claimed in claim 3, in which the wall extends generally parallel to the heat transfer surface or the surface of the component.

5. An installation as claimed in claim 3, in which the wall is part of a partition which divides the chamber into upstream and downstream compartments, with respect to the direction of fluid flow through the chamber.

6. An installation as claimed in claim 3, in which the flow passage is one of an array of flow passages in the wall.

7. An installation as claimed in claim 3, in which the or each flow passage comprises a hole in the wall.

8. An installation as claimed in claim 7, in which the wall material defining the or each flow passage is inclined to the plane of the wall, thereby to direct fluid passing through the respective flow passage.

9. An installation as claimed in claim 3, in which the wall is provided with louvers, and the or each flow passage comprises a channel defined between adjacent louvers.

10. An installation as claimed in claim 9, in which the or each channel has an inlet aligned substantially with the general direction of fluid flow through the chamber, and an outlet directed towards the heat transfer surface or the surface of the component.

11. An installation as claimed in claim 2, in which the component is a combustion engine, the chamber being defined between the engine and an engine enclosure.

12. An installation as claimed in claim 11, in which the engine is a gas turbine engine and the enclosure is a nacelle, the chamber being annular, and defined between the nacelle and the engine.

13. An installation as claimed in claim 12, when appendant to claim 3, in which the wall is part of a partition extending between the engine and the nacelle.

14. An installation as claimed in claim 11, in which the surface to be cooled is a turbine casing of the engine.

Description:

This invention relates to an installation having a thermal transfer arrangement, and is particularly, although not exclusively, concerned with an installation comprising a component, such as a gas turbine engine, and a cooling arrangement for cooling the component.

Gas turbine engines generate very high temperatures in operation, and it is usual for internal cooling arrangements to be provided for cooling parts such as turbine blades and vanes. It is also known for gas turbine engine installations to provide cooling arrangements for cooling external surfaces of the engine, for example regions of the turbine casing. For example, U.S. Pat. No. 5,100,291 discloses a cooling arrangement which comprises spray bars surrounding regions of the engine casing to be cooled. Air is supplied to the spray bars from the compressor of the engine, and emerges from the spray bars through openings which direct air towards the casing to provide impingement cooling.

Air taken from the compressor for cooling purposes is removed from the main gas flow path through the engine, and so does not contribute to the power generated by the engine. Consequently, the efficiency of the engine is reduced, and in particular the specific fuel consumption (SFC) is increased, for example by approximately 0.7% for each kilogram of coolant air extracted from the compressor. Furthermore, air extracted from high pressure compressor stages may be hot, so reducing its cooling effectiveness.

U.S. Pat. No. 6,227,800 discloses an alternative cooling arrangement which avoids use of high pressure air extracted from the compressor of the engine. In the arrangement of U.S. Pat. No. 6,227,800, the engine is situated within a bay defined on the outside by a nacelle. The bay is ventilated by a flow of air generated by a fan of the engine. The air flows over the engine casing and is constrained, in the region of the turbine of the engine, to flow through a narrow passage defined by an annular baffle. The air is accelerated through the passage to cool the turbine casing by forced convection.

The narrow passage thus enhances the cooling effect of the general air flow through the bay, parallel to the surface of the engine casing. However, effective cooling of the turbine casing using a cooling arrangement as disclosed in U.S. Pat. No. 6,227,800 requires a substantial mass flow rate of air and a substantial driving pressure to force the air through the narrow passage. The required mass flow rate and pressure are substantially greater than are typically found in gas turbine engine bays. Furthermore, as the cooling air travels through the narrow passage, it picks up heat so that the cooling effectiveness diminishes in the downstream direction, with the result that regions of the turbine casing close to the outlet of the narrow passage may be inadequately cooled.

There is consequently a need for a cooling arrangement which can provide adequate cooling of a gas turbine engine casing while utilising relatively low pressure air available in the engine bay.

According to the present invention there is provided an installation comprising a component having a means for inducing fluid flow through the component, and a thermal transfer arrangement, the thermal transfer arrangement comprising a chamber which is ventilated in operation of the installation by a forced fluid flow induced by a means different to the means for inducing fluid flow through the component, the chamber being bounded at least partly by a heat transfer surface of the component, a flow passage being disposed within the chamber which is oriented to direct at least part of the fluid flow at the heat transfer surface, to transfer heat between the fluid and the heat transfer surface.

A more specific aspect of the present invention provides an installation comprising a component having a means for inducing fluid flow through the component, and a cooling arrangement for cooling the component, the cooling arrangement comprising a chamber which is ventilated in operation of the installation by a forced fluid flow induced by a means different to the means for inducing fluid flow through the component, the chamber being bounded at least partly by a surface of the component, a flow passage being disposed within the chamber, which flow passage is oriented to direct at least part of the fluid flow at the surface of the component to cool the component.

The flow passage may be provided in a wall disposed within the chamber. The wall may be generally parallel to the heat transfer surface or the surface of the component, and may comprise part of a partition which extends across the chamber to divide the chamber into upstream and downstream compartments, with respect to the direction of fluid flow through the chamber in operation.

The flow passage may be one of an array of flow passages provided in the wall. The or each flow passage may comprise a hole in the wall, in which case the wall material defining the or each flow passage may be inclined to the plane of the wall, thereby to direct fluid flow at the heat transfer surface or the surface of the component.

Alternatively, or in addition, the wall may be provided with louvers, so that the or each flow passage comprises a channel defined between adjacent louvers. The louvers may be configured so that the or each channel has an inlet substantially aligned with the general direction of fluid flow through the chamber, and an outlet directed towards the heat transfer surface or the surface of the component.

If the installation comprises a component to be cooled, the component may be a combustion engine, and the chamber may be defined between the engine and an engine enclosure. The engine may be a gas turbine engine, and the enclosure may be a nacelle, with the result that the chamber is annular, defined between the nacelle and the engine. Where the or each flow passage is provided in a partition, the partition may extend between the engine and the nacelle.

The surface to be cooled may be a turbine casing of the engine.

For a better understanding of the present invention, and to show more clearly how it may be carried into effect, reference will now be made, by way of example, to the accompanying drawings, in which:

FIG. 1 is a schematic sectional view of a gas turbine engine installation;

FIG. 2 is an enlarged view of part of the installation of FIG. 1;

FIGS. 3 to 8 show variants of a cooling arrangement in the installation shown in FIGS. 1 and 2;

FIG. 9 is an enlarged sectional view corresponding to the cooling arrangements shown in FIGS. 3 to 7; and

FIGS. 10 and 11 are sectional views of the cooling arrangement shown in FIG. 8.

FIG. 1 shows an engine installation comprising a gas turbine engine 2 enclosed by a nacelle 4. The engine 2 is axisymmetric about an axis X.

As is conventional, the engine 2 comprises a compressor 6, a combustor 8 and a turbine 10. The components 6, 8 and 10 of the engine 2 are enclosed in an engine casing 12. The casing 12 is assembled from a series of axially interconnected annular components and sub-assemblies, and in particular includes a turbine casing 14.

An annular engine bay 16 is defined between an outer wall 18 of the nacelle 4 and the engine casing 12. Various ancillary components of the engine 2 are accommodated within the bay 16, such as an oil cooler unit 20, pipework 22 and other ancillaries which are not shown but which may be, for example, gearboxes, air pipes and electronic packages such as health monitoring and control boxes. The bay 16 is closed at its axial ends, with respect to the axis X, by a forward bulkhead 24 and an aft bulkhead 26. An air intake 28 opens at the forward end of the nacelle 4 and extends through the forward bulkhead 24 into the bay 16. A bay exhaust (not shown) is provided at or close to the aft bulkhead 26.

In the region of the bay 16 surrounding the turbine 10, there is a partition 28 which extends between the turbine casing 14 and the wall 18 of the nacelle 4, and divides the bay 16 into upstream and downstream compartments 40 and 42. The partition 28 is shown in more detail in FIG. 2, and comprises an outer solid wall 30 extending radially inwardly from the nacelle wall 18, an inner solid wall 32 extending obliquely radially outwardly and in the aft direction from the turbine casing 14 and a permeable wall 34 extending between the radially inner end of the outer solid wall 30 and the radially outer end of the inner solid wall 32.

As shown diagrammatically in FIG. 2, the permeable wall 34 is provided with an array of flow passages 36 which are oriented generally towards the turbine casing 14. For this purpose, the permeable wall 34 extends generally parallel to the region of the turbine casing 14 that it surrounds.

In operation during flight, movement through the air of the aircraft in which the engine 2 is installed will cause air to enter the bay 16 through the intake 27. This air travels through the bay 16 in the general direction indicated by arrows 38 to ventilate the bay 16 to remove fumes and vapour and to cool ancillary components within the bay 16, such as the oil cooler unit 20 and the pipework 22. Typically, the flow rate of the air through the bay 16 is sufficient for five volume changes per minute. The air flow is deflected by the partition 28 and flows through the flow passages 36, as will be described in more detail below, to enter the aft compartment 42. The air then exits the bay 16 through the bay exhaust, which may take the form of discrete holes in the aft bulkhead 26 which discharge into the ambient surroundings at the aft of the nacelle 4. Alternatively, the bay exhaust may comprise a network of passages conveying the air away from the compartment 42 in a manner which minimises cross flow interference.

It will be appreciated that the air deflected by the partition 28 is directed towards the turbine casing 14, as indicated by an arrow 44, and thus flows through the flow passages 36 to emerge as a series of jets indicated by arrows 46. These jets 46 impinge on the turbine casing 14 in a direction substantially normal to the surface of the turbine casing 14, to achieve an impingement cooling effect.

While it is not essential for the flow passages 36 to direct the jets 46 normal to the surface of the turbine casing 14, this will achieve an optimum impingement cooling effect. However, adequate impingement cooling can be achieved with the jets 46 directed at angles of less than 90° to the surface of the turbine casing 14. For example, while it is desirable for the angle of impingement to be greater than 40°, lower impingement angles can nevertheless provide sufficient cooling in some circumstances. As a general rule, suitable angles of impingement will be greater than 60° and preferably greater than 75°. It will be appreciated that, as a consequence it is not essential for the permeable wall 34 to be precisely parallel to the turbine casing 14 so long as the jets 46 are oriented so as to provide adequate impingement cooling.

Following impingement of each jet of air 46 against the turbine casing 14, the air will “rebound” as a fountain of air directed away from the surface of the turbine casing 14, and will then flow in the direction of the arrow 48 through the compartment 42 to the bay exhaust.

A pressure control screen 47 may be provided which extends between the turbine casing 14 and a point on the partition 28 at or close to the junction between the outer solid wall 30 and the permeable wall 34. The screen 47 serves to control the pressure drops between the upstream compartment 40 and the region 49 within the permeable wall 34, and between that region 49 and the rest of the downstream compartment 42.

It will be appreciated that the jets of air 46 will be at substantially the same temperature as each other when they impinge on the turbine casing 14. Consequently, the turbine casing 14 can be adequately cooled over substantially its full extent. In some circumstances, it may be desirable to vary the extent of cooling across the surface of the turbine casing 14, in order to provide additional cooling to areas that have a greater cooling requirement. This can be achieved by suitably arranging the flow passages 36, both in terms of their position and their orientations. For example, it has been found that increased cooling may be required in regions of the turbine casing 14 where hooks are provided to attach internal liners to the casing 14.

Additional measures may be provided to avoid hot spots in the region 49, such as arise from stagnant flow. For example, there may be provision for air flow through the inner solid wall 32.

In a typical gas turbine engine installation in an aircraft, the nominal mass flow rate of air through the bay 16 at maximum take-off (MTO) thrust, when the requirement for cooling is greatest, is less than 0.5 Kg/s. This is substantially below the flow rate required to achieve adequate cooling by means of a cooling arrangement as disclosed in U.S. Pat. No. 6,227,800, where the air flow is constrained to pass through a narrow gap while travelling along the surface to be cooled. It is estimated that, by employing impingement cooling in the manner described above with reference to FIGS. 1 and 2, the heat transfer coefficient between the air flowing through the bay 16 and the turbine casing 14 can be improved by a factor of 4 over the cooling arrangement disclosed in U.S. Pat. No. 6,227,800. The increased cooling effect by the air flowing in the bay 16 reduces the requirement for internal cooling, i.e. cooling of internal components of the engine using air bled from the compressor 6. This saving of compressor air used for cooling results in an improvement in the SFC of the engine, leading to a reduction, of the order of 0.27%, in the fuel burn of the engine.

This flow passages 36 may be formed in a variety of ways, and may differ across the extent of the permeable wall 34. FIGS. 3 to 11 show various forms of flow passages 36, although it will be appreciated that other variants are possible.

FIG. 3 shows flow passages 36 in the form of circular holes which are distributed over the permeable wall 34 in a regular manner, both in the circumferential direction and the axial direction. By contrast, FIG. 4 shows a more random array of holes 36, in which the holes in each circumferential row are offset with respect to one another. That is to say, the circumferential rows of holes may be staggered relative to one another. Each circumferential row of holes may be offset mid-way circumferentially between the neighbouring rows of holes.

FIG. 5 illustrates an embodiment in which the holes 36 vary in size over the extent of the permeable wall 34, while FIG. 6 illustrates an embodiment in which the holes 36 are non-circular. Another example of non-circular holes 36 is shown in FIG. 7, where the holes 36 are in the form of slots. In FIG. 7, the slots are illustrated as aligned with the circumferential direction, but it will be appreciated that other orientations of the slots 36 are possible.

The holes and slots 36 shown in FIGS. 3 to 7 may be plain holes, but there may be advantages in profiling the edges of the holes in order to minimise the pressure drop across them, and to direct the air flowing through them in the desired direction. For example, the holes 36 may be plunged, in other words the material of the permeable wall 34 surrounding the holes may be preformed out of the plane of the wall 34, for example in the downstream direction with respect to the flow direction through the holes 36. Alternatively, as shown in FIG. 9, different regions of the periphery of each hole 36 may be deformed in different directions. As shown in FIG. 9, the hole or slot 36 is “turned” in a “cheese-grater” manner, i.e. the material of the wall 34 surrounding the hole or slot is deflected from the plane of the wall 34 in opposite directions on opposite sides of the hole or slot 36, as indicated by the deflected regions 48, 50. These deflected regions create a duct extending through the permeable wall 34, to direct the air in the manner indicated by arrows 52.

FIG. 8 is a view corresponding to those of FIGS. 3 to 7, but illustrates an embodiment in which the permeable wall 34 is provided with an array of circumferentially extending louvers 54. The flow passages 36 are thus defined by the spaces between adjacent louvers 54. As will be appreciated, it is not essential for the louvers to extend circumferentially; they could instead extend generally axially, or in an oblique direction.

As shown in FIG. 10, the louvers 54 may be flat (i.e. planar) in which case the flow passages 36 are in the form of straight parallel-sided ducts. Alternatively, the louvers may be curved, or otherwise profiled, to provide flow passages 36 in the form of curved ducts which serve to deflect the general air flow 38 from a generally axially direction to a direction 56 directed towards the surface of the turbine casing 14 (not shown in FIGS. 10 and 11).

The louvers may be adapted, for example by additional circumferential shaping, to enhance the impingement cooling effect.

Additionally, the external surface of the turbine casing 14 may be provided with features, such as pedestals, to improve heat transfer between the impinging air and the turbine casing 14.

The invention has been described with reference to FIGS. 1 and 2 in the context of a gas turbine engine installation in an aircraft, in which the air flow in the bay 16 is generated by the forward motion of the aircraft through the air. However, it will be appreciated that the cooling arrangement described could be employed in installations, such as stationary installations, in which the air flow in the bay 16 is generated by other means, for example a fan. A cooling arrangement in accordance with the present invention can be employed for cooling engines other than gas turbine engines, for example reciprocating engines and steam engines.

Furthermore, a cooling arrangement in accordance with the present invention can be employed for cooling components other than engines, where a significant heat transfer coefficient is required but only a low pressure drop is available to generate a flow of cooling air. For example, a cooling arrangement in accordance with the present invention could be employed to cool electronic components, such as computer chips.

While the invention has been described with reference to the cooling of components, it will be appreciated that embodiments in accordance with the present invention may be employed to heat components using a flow of hot fluid driven by a relatively low pressure drop.

In the embodiment described with reference to FIGS. 1 and 2, the region 49 is a continuous annular region extending around the turbine casing, which communicates with the rest of the downstream compartment 42 either directly or through the screen 47. In an alternative embodiment, the region 49 could comprise the interior of a large ring installed around the turbine casing 14. The region 49 need not be continuous, but could be segmented to permit axial and/or circumferential air distribution control.

Embodiments in accordance with the present invention provide cooling, in particular of a turbine casing, at low or no cost while achieving high heat transfer rates. The present invention can be put into practice in a simple manner with little weight penalty and at a low manufacturing cost. In embodiments in accordance with the invention, there is little or not interference with surrounding components and structures in the engine bay 16, and the features of the invention can be integrated easily with the rest of the engine design.