INSTALLATION HAVING A THERMAL TRANSFER ARRANGEMENT
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.
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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:
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
As shown diagrammatically in
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
This flow passages 36 may be formed in a variety of ways, and may differ across the extent of the permeable wall 34.
The holes and slots 36 shown in
As shown in
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
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
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.
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.
Type: Application
Filed: Apr 7, 2011
Publication Date: Oct 27, 2011
Applicant: ROLLS-ROYCE PLC (London)
Inventor: Christopher S. AVENELL (Bristol)
Application Number: 13/081,938
International Classification: F01D 25/12 (20060101); F28D 20/00 (20060101);