WO2004057157A1

WO2004057157A1 - Turbine blade

Info

Publication number: WO2004057157A1
Application number: PCT/GB2003/005606
Authority: WO
Inventors: John Macdonald; Kevin Dorling
Original assignee: John Macdonald; Kevin Dorling
Priority date: 2002-12-21
Filing date: 2003-12-19
Publication date: 2004-07-08
Also published as: GB0229908D0; AU2003295141A1

Abstract

A turbine blade (10) is described which includes an aerofoil section (12) mounted on a blade root (18), wherein a plurality of internal cooling channels (24) extend through the blade root (18) and aerofoil section (12) to provide a flow passage for a cooling medium. The cooling channels (24) include a plurality of axially spaced and inwardly projecting turbulence promoting ribs (30) located on a wall surface (32) of the channels (24). The ribs (30) are provided along the entire length of the portion of the channels (24) which extend through the aerofoil section (12), while the portion of the channels (24) which extend through the blade root (18) do not include ribs (24), and in one disclosed embodiment, the terminating ends of the channels (24) at the tip (14) of the aerofoil section (12) include a rib (30). Methods of manufacturing the internal cooling channels in the blade are also disclosed and comprise an electrochemical machining method. In this method, an electrode is used to penetrate the blade. By varying the rate of advancement of the electrode, smaller and larger inner bore diameters are produced and result in a cooling channel with a plurality of ribs.

Description

TURBINE BLADE

The present invention relates to turbomachinery, and in particular, but not exclusively, to turbine blades for use in gas turbine engines .

Gas turbine engines are used in a number of applications, including aircraft propulsion systems and power generation systems and the like. Typical gas turbine engines generally consist of three components: a compressor, a combustion chamber, and a turbine unit, wherein the compressor and the turbine unit are mounted on the same shaft. In use, air is compressed by the compressor, is fed into the combustion chamber where it is mixed with fuel and the mixture is ignited, and the exhaust gases produced are then expanded through the turbine unit to drive the shaft and produce shaft work. In power generation applications, the shaft work produced is used to drive the compressor and turn electrical generators, often via a gearing system. Conventional turbine units comprise a plurality of stages, each stage usually consisting of two sets of blades arranged in an annulus, the first set being stator or nozzle blades which are rotationally fixed with respect to the casing of the turbine, and the second set being rotor blades which are mounted on the shaft and rotate therewith. The number of stages in a turbine unit is selected in accordance with, for example, considerations of stage mechanical loading and thermodynamic performance. Additionally, the number of stages may be determined by the required pressure ratio from turbine inlet to outlet .

Turbine efficiency is an important factor in the design of any gas turbine engine and one method of increasing the performance characteristics involves maximising the temperature of the gas at the turbine inlet. However, increasing the temperature of the gas used to drive the turbine produces serious mechanical and thermal stressing problems in the turbine blades, and the temperature of the gas is limited by the physical properties of the blade material, such as melting point and yield strength and the like.

Various advancements in materials have been made for use in high pressure and temperature turbines, however, these materials are extremely costly due to the complex formation process, for example, such as uni-directional crystallisation.

It is therefore common practice to minimise the thermal stress by cooling the blades during operation by passing cooling air bled from the compressor externally and internally of the blades, such that higher operational temperatures may be achieved, and the life span of the blades may be increased. A number of blade designs currently exist which allow a particular cooling air flow regime to be utilised to allow a combination of, for example, convection cooling, impingement cooling and film cooling in order to improve the heat transfer properties between the blade and the cooling air. However, the actual shape or design of a blade is often determined by a compromise between aerodynamic and integrity requirements. Cooling primarily affects the integrity considerations both in terms of controlling the thermal stresses and maintaining the operating temperature of the material within acceptable limits to minimise creep and corrosion.

It is among the objects of the present invention to provide a turbine blade having improved cooling.

According to a first aspect of the present invention, there is provided a turbine blade including an aerofoil section comprising opposing pressure and suction side walls adjoining at leading and trailing edges of the blade, and defining at least one internal cooling channel extending through the aerofoil section from a base to a tip thereof, said at least one channel including a plurality of axially spaced and inwardly projecting ribs located on a wall surface of the channel, said ribs being turbulence promoting ribs and being provided along the entire length of the at least one channel. Thus, in use, the at least one internal cooling channel allows a cooling medium such as compressed air to be passed therethrough to provide internal cooling of the blade, and wherein the ribs induce turbulence in the cooling medium resulting in an increased cooling effect by increasing the effective heat transfer coefficient of the cooling medium.

The turbine blade of the present invention provides for improved heat transfer between the blade and a cooling medium over the entire length of the aerofoil section due to the presence and form of the ribs provided within the at least one channel extending therethrough. For example, the provision of turbulence promoting ribs in the region of the tip of the aerofoil section results in a reduced operating temperature in this region which may include a tip shroud which is susceptible to high temperature creep failure.

Preferably, the turbine blade further includes a root portion, wherein the root portion includes a blade platform, upon which blade platform the aerofoil section is mounted via its base. In one embodiment, the aerofoil section and root portion may be integrally formed. Alternatively, the aerofoil section and root portion may be formed separately and subsequently secured together, for example, by welding or the like. Advantageously, the root portion defines at least one internal cooling channel substantially aligned with the at least one cooling channel extending through the aerofoil section. Preferably, the root portion defines a 5 corresponding number of cooling channels as are defined in the aerofoil section. In a preferred embodiment of the present invention, the at least one channel of the root portion is smooth and does not include any turbulence promoting ribs. Alternatively, the at least

10 one channel of the root portion may include one or more axially spaced turbulence promoting ribs. Conveniently, in use, a cooling medium may be communicated from the at least one channel of the root portion and into an associated at least one channel of the aerofoil section.

15 Preferably, the at least one channel of the root portion and the at least one channel of the aerofoil section together define an at least one blade cooling channel which extends through the root portion, through the aerofoil section and terminates at the tip of the

20. aerofoil section. Advantageously, the at least one blade cooling channel is open at the root portion and tip of the aerofoil section, thus forming respective apertures at the root and tip of the blade. Consequently, a cooling medium may be supplied to the blade via the

25 respective aperture in the root portion and exit the blade via the respective aperture in the tip of the aerofoil section.

Advantageously, the at least one blade cooling channel is substantially aligned with the general

30 longitudinal axis of the turbine blade.

Conveniently, the at least one blade cooling channel displays a change in orientation at an interface between the root portion and the aerofoil section of the blade. That is, the longitudinal axis of the at least one channel of the root portion is angularly offset from the longitudinal axis of a corresponding at least one channel of the aerofoil section.

In a preferred embodiment of the present invention, the at least one channel of the aerofoil section has a substantially circular lateral cross-sectional shape. The at least one channel of the aerofoil section may alternatively have a substantially square, rectangular or triangular cross-section, or any other suitable cross- sectional shape.

Advantageously, the at least one channel of the root portion may have the same general lateral cross-sectional shape as the at least one channel of the aerofoil section. Alternatively, the at least one channel of the root portion may have a different lateral cross-section shape from the at least one channel of the aerofoil section, as required.

Preferably, where the at least one channel of the aerofoil section is of circular cross-section, the turbulence promoting ribs are annular ribs which extend radially inwardly from the wall surface of the channel so as to define annular recesses between adjacent ribs. Preferably also, the turbulence promoting ribs each extend uninterrupted around the inner perimeter of the at least one channel. Alternatively, at least one of the ribs may include at least one gap therein.

Conveniently, the ribs may extend substantially perpendicular from the surface of the at least one channel of the aerofoil section. Alternatively, the ribs may extend at an acute angle from the surface of the at least one channel of the aerofoil section with respect to the flow direction such that each rib is directed into or against the flow direction of cooling medium. Alternatively further, the ribs may extend at an obtuse angle from the surface of the at least one channel of the aerofoil section with respect to the flow direction such that each rib is directed in the same direction as the flow of cooling medium. The ribs may, however, extend at any angle or combination of angles noted above.

Preferably, the distribution of the turbulence promoting ribs is such that the terminating end of the at least one channel of the aerofoil section at the tip thereof includes a turbulence promoting rib. In a preferred embodiment of the present invention, the general dimension of the at least one channel of the aerofoil section normal to the intended direction of flow of cooling medium, for example, the diameter or width of the channel, varies along the length thereof. Advantageously, the diameter or width of the at least one channel in the region of the base of the aerofoil is in the range of 3 to 5 mm, and preferably between 3.4 to 4.6 mm, and the diameter or width in the region of the tip of the aerofoil section is in the region of 3 to 5 mm, and preferably between 3.4 to 4.6 mm. Preferably, the diameter of the at least one channel in a central region of the aerofoil section of the blade is around 3 to 5.2 mm, and more preferably between 3.4 to 5.1 mm.

Preferably, the diameter or width of the at least one channel of the root portion is uniform along the length thereof and is preferably in the region of 6 to 7 mm, and more preferably around 6.7 mm.

Preferably also, the distance between adjacent ribs, generally termed the rib pitch, varies over the length of the at least one channel of the aerofoil section. Additionally, or alternatively, the rib height may vary over the length of the at least one channel .

Advantageously, the rib height is greatest and the rib pitch is smallest in the central region of the aerofoil section of the turbine blade. This arrangement maximises the heat transfer between the blade and the cooling medium in this region by the provision of ribs which are more densely packed and which impinge into the flow of cooling medium to a greater extent. In one disclosed embodiment of the present invention, the rib pitch in the central region of the aerofoil section may range between 3 and 5.5 mm, and the rib height in this region may range between 0.15 and 0.39 mm. Advantageously, the rib height may be smallest and the rib pitch may be greatest in cooler regions which require a lower rate of heat transfer, such as in the regions of the base and tip of the aerofoil section of the blade, which are generally cooler during operation in relation to, for example, the central region. The rib height and pitch in the base and tip regions of the aerofoil section may be, for example, around 0.15mm and 7.5 mm respectively.

Preferably, the aerofoil section of the turbine blade defines a plurality of internal cooling channels distributed in a general chord-wise direction of the blade between the leading and trailing edges of the aerofoil section. Conveniently, the root portion of the blade defines a corresponding number of internal cooling channels arranged to be aligned with a respective cooling channel of the aerofoil section. In one disclosed embodiment of the present invention, eight cooling channels are provided. However, it should be understood that any number of channels may be utilised, as required, in accordance with considerations such as blade dimensions and cooling requirements and the like.

Conveniently, where a plurality of internal cooling channels is provided, the various properties and features of any turbulence promoting ribs may vary from channel- to-channel. Thus, by varying the dimensions and features of the channels and the ribs in the manner described above allows the optimum heat transfer between the cooling medium and the blade to be achieved at specific points of the blade.

In a preferred embodiment of the present invention, the width of the ribs is constant over the entire length of the at least one channel of the aerofoil section. Preferably, the ribs are each approximately 1 mm wide. Alternatively, the rib widths may vary along the channel length, and/or, where a plurality of channels is provided, from channel-to-channel.

Advantageously, the root portion of the blade may be of a fir-tree type. Alternatively, the root portion may be of a dove-tail type, or any other type commonly used in the art .

Preferably, the cooling medium is air, and more preferably air supplied from a compressor.

Advantageously, the turbine blade may be for use in a gas turbine engine .

Preferably, the turbine blade is a rotor blade. Alternatively, the turbine blade may be a stator or nozzle blade.

More preferably, the turbine blade is a second stage rotor blade for use in a gas turbine engine.

According to a second aspect of the present invention, there is provided a gas turbine engine including a plurality of turbine blades, at least one turbine blade including an aerofoil section comprising opposing pressure and suction side walls adjoining at leading and trailing edges of the blade, and defining at least one internal cooling channel extending through the aerofoil section from a base to a tip thereof, said at least one channel including a plurality of axially spaced and inwardly projecting ribs located on a wall surface of the channel, said ribs being turbulence promoting ribs and being provided along the entire length of the at least one channel . According to a third aspect of the present invention, there is provided electrical generating means including a gas turbine engine, said gas turbine engine including a plurality of turbine blades, at least one turbine blade including an aerofoil section comprising opposing pressure and suction side walls adjoining at leading and trailing edges of the blade, and defining at least one internal cooling channel extending through the aerofoil section from a base to a tip thereof, said at least one channel including a plurality of axially spaced and inwardly projecting ribs located on a wall surface of the channel, said ribs being turbulence promoting ribs and being provided along the entire length of the at least one channel .

Preferably, the electrical generating means is adapted for use in a power producing plant such as a fossil fuel power plant or a nuclear power plant or the like.

According to a fourth aspect of the present invention, there is provided a method of forming at least one cooling channel in a turbine blade by an electrochemical machining operation, said method comprising the steps of :

(a) providing a turbine blade having an aerofoil section mounted on a root portion; (b) introducing an electrode to at least one of the root portion and a tip of the aerofoil section;

(c) activating and advancing said electrode to penetrate the blade to form a first bore; and (d) varying the rate of advancement of the electrode such that an increased rate of advancement will produce a smaller inner bore diameter and a reduced rate of advancement will produce a larger inner bore diameter, such that a plurality of upstanding ribs are formed between adjacent recesses in the first bore; such that at least one cooling channel is formed in the blade which includes a plurality of upstanding annular ribs along the entire length of said aerofoil section.

In one embodiment of the present invention the electrode is advanced in the manner described in step (d) above until the electrode emerges from the other of said root portion and tip of the aerofoil section. Alternatively, the method may involve the additional steps of :

(e) introducing an electrode to the other of said root position and tip of the aerofoil section;

(f) activating and advancing the electrode to penetrate the blade to form a second bore; and

(g) advancing the electrode until the second bore merges with the first bore.

In one embodiment, the first and second bores may be formed simultaneously. Alternatively, the second bore may be formed before or after the first bore is formed.

Conveniently, the first and second bores may be formed to be angularly off-set.

Advantageously, the electrode used to form the second bore may be advanced at a rate selected in accordance with the required bore wall structure. For example, where ribs are required, the rate of advancement of the electrode forming the second bore may be varied. Alternatively, where no ribs are required, the electrode may be advanced at a constant rate. Preferably, the method of the present invention is adapted to form a smooth bore in the rib portion of the blade. It should be understood that the term smooth means that the channel of the root portion does not include any ribs .

According to a fifth aspect of the present invention, there is provided a method of manufacturing a turbine blade, said method comprising the steps of: forming a blade having an aerofoil section mounted on a root portion; forming at least one cooling channel in said blade by an electrochemical machining process wherein;

(a) an electrode is introduced to at least one of the root portion and a tip of the aerofoil section;

(b) said electrode is activated and advanced to penetrate the blade to form a first bore; and

(c) the rate of advancement of the electrode is varied such that an increased rate of advancement will produce a smaller inner bore diameter and a reduced rate of advancement will produce a larger inner bore diameter, such that a plurality of upstanding ribs are formed between adjacent recesses in the first bore; such that at least one cooling channel is formed in the blade which includes a plurality of upstanding annular ribs along the entire length of the aerofoil section.

Preferably, the method of the present invention is adapted to form a smooth bore in the root portion of the blade. It should be understood that the term smooth means that the channel of the root portion does not include any ribs.

In one embodiment of the present invention the electrode is advanced in the manner described in step (c) above until the electrode emerges from the other of said root portion and tip of the aerofoil section.

Alternatively, . the method may involve the additional steps of : (d) introducing an electrode to the other of said root position and tip of the aerofoil section;

(e) activating and advancing the electrode to penetrate the blade to form a second bore; and

(f) advancing the electrode until the second bore merges with the first bore.

In one embodiment of the present invention, the first and second bores may be formed simultaneously. Alternatively, the second bore may be formed before or after the first bore is formed. Conveniently, the first and second bores may be formed to be angularly off-set.

Advantageously, the variation of the rate of advancement of the electrode may be such that the width of the recesses varies along the length of the at least one channel.

The rate of advancement of the electrode may also be varied in such a manner that the subsequent height of each rib varies along the length of the at least one channel . Preferably, the rate of advancement of the electrode in the region of the tip of the aerofoil section is such that an upstanding rib is formed at the terminating end of the at least one channel .

Preferably also, the electrochemical machining process is repeated such that a plurality of cooling channels may be formed.

Conveniently, multiple bores may be created simultaneously using a comb of electrodes in which multiple electrodes are arranged to be activated and advanced at the same time to penetrate the blade and form a plurality of bores.

These and other aspects of the present invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

Figure 1 is a diagrammatic representation of a turbine blade in accordance with an embodiment of the present invention;

Figure 2 is an enlarged diagrammatic representation of a portion of a cooling channel of the blade of Figure l;

Figure 3 is another diagrammatic representation of the turbine blade of Figure 1, showing different regions of interest, A-E; Figure 4 is an enlarged diagrammatic representation of a portion of a cooling channel in region A, as shown in Figure 3 ;

Figure 5 is an enlarged diagrammatic representation of a portion of a cooling channel in regions B-E, as shown in Figure 3;

Figure 6 shows a detailed view of a cooling channel in the region of the tip of the blade of Figure 1;

Figure 7 shows a graphical representation of an electrochemical process for forming a cooling channel in the blade of Figure 1; and

Figure 8 is a general diagrammatic illustration of the form of a channel created by the process shown in the graphical representation of Figure 7.

Reference is first made to Figure 1 of the drawings in which there is shown a turbine blade, generally indicated by reference numeral 10, for use in a gas turbine engine in accordance with an embodiment of the present invention. The blade 10 is a second stage rotor blade and includes an aerofoil section 12 having a tip 14 and a base 16, wherein the aerofoil section 12 is mounted on a blade root 18. The aerofoil section 12 has opposing pressure and suction side walls adjoining at a leading edge 20 and a trailing edge 22 of the blade 10. Although not shown, the blade 10 may include a blade shroud located on the tip 14. The turbine blade 10 defines a number of internal cooling channels 24, which channels are circular in cross-section and provide a flow passage for a cooling medium, such as compressed air, to cool the blade while in use. Although three channels 24 are shown in Figure 1, this is only representative for the purpose of improving clarity of the drawings, and it should be understood that any number of channels may be present.

The channels 24 are of a single-pass type and extend through the blade root 18 and through the aerofoil section 12 of blade 10. In the embodiment shown, the portion of the channel extending through the blade root 18 is angularly off-set from the portion of the channel extending through the aerofoil section 12. In use, a cooling medium is supplied to the internal channels 24 through apertures 26 in the blade root 18, and exits the channels 24 through respective apertures 28 in the tip 14 of the aerofoil section.

As shown in Figure 2, which is an enlarged diagrammatic representation of a portion of a channel 24 of the blade 10, the channels 24 include a plurality of turbulence promoting ribs 30 provided along the length of the channel wall surface 32, wherein the ribs 30 seek to improve the heat transfer between the surfaces of the blade 10 and the cooling medium. As shown, the ribs 30 are upstanding from the wall surface 32 of the channel 24 so that they produce axially spaced restrictions extending into the flow path defined by the channel 24, and hence create a turbulent flow regime. In the embodiment shown, the ribs 30 are uninterrupted and annular and thus define annular recesses 34 between adjacent ribs 30. Additionally, the ribs 30 extend perpendicularly from the surface 32 of the channel 24. It is a feature of the present invention that ribs

30 are provided over the entire length of those portions of the channels 24 which extend through the aerofoil section 12. This arrangement maximises the overall heat transfer between the blade 10 and cooling medium which generates a number of advantages. For example, providing ribs 30 at the tip 14 of the blade 10 results in a reduced operating temperature in this region which may include a blade shroud which is susceptible to high temperature creep failure. As generally represented in Figure 2, the distribution of the turbulence promoting ribs 30 is such that the rib pitch varies along the length of the channel 24, the rib pitch being understood to be the distance between adjacent ribs 30. Additionally, the rib height and channel diameter vary over the length of the channel, as discussed below with reference to Figures 3 to 5 and Table 1.

Referring first of all to- Figure 3, there is shown another diagrammatic representation of the turbine blade 10 of Figure 1, wherein different regions of interest along the length of the blade 10 are shown, indicated by letters A-E, wherein: 'A' represents the region of the root 18 of the blade 40; 'B' represents the region of the base 16 of the aerofoil section 12; ' C represents a lower-middle region of the aerofoil section 12; 'D' represents and upper-middle region of the aerofoil section 12; and 'E' represents the region of the tip 14 of the aerofoil section 12. In the present embodiment, eight internal cooling channels 24 are provided, however, for clarity, only two channels 24 are shown in Figure 3, one in the region of the leading edge 20 of the blade 10, and one in the region of the trailing edge 22 of the blade 10. For the purposes of the following description, the channels of the blade 10 of Figure 3 are numbered 1 through 8, with channel number 1 being located in the region of the leading edge 20, and channel number 8 being located in the region of the trailing edge 22, with channel numbers 2 to 7 being distributed in order between channels 1 and 8.

As noted above with reference to Figure 2, the rib height, rib pitch and channel diameter vary over the length of the channel. In addition to this, the rib height, rib pitch and channel diameter varies from channel-to-channel . However, it should be noted that the portion of the channels 24 extending through region A, which have a length of approximately 130 mm, do not include any ribs, as shown in Figure 4, and define a substantially uniform diameter 40 of 6.7 mm. It is possible to eliminate the ribs from the channels of region A due to the fact that this region requires the lowest rate of heat transfer.

The specific variations of the rib height (represented by dimension X in Figure 5) , rib pitch (represented by dimension Y in Figure 5) , and channel diameter (represented by dimension Z in Figure 5) , of the channels (numbers 1 through 8) through regions B-E are outlined in Table 1 below.

TABLE 1

It should be noted that channel number 7 has been omitted from Table 1. In this regard, the characteristics of channel number 7 may be the same as those of either channel 6 or channel 8, or alternatively may define relevant characteristics in a range between those of channels 6 and 8. Referring to Table 1, it is apparent that in the middle region of the blade 10, that is, regions C and D, the rib height of the channels is greatest and the rib pitch is smallest in comparison to the other regions. This arrangement maximises the heat transfer between the blade 10 and the cooling medium in this region by the provision of ribs 30 which are more densely packed and which impinge into the flow of cooling medium to a greater extent.

Table 1 also shows that the diameter of the channel numbers 1 to 5 in each region B-E is greater than that of the remaining channels 6 to 8. This is due to the fact that the leading edge portion of the blade is generally hotter than the trailing portion and as such requires a larger rate of cooling, which is provided for by including larger channels allowing for an increased mass flow rate of cooling medium.

Varying the dimensions of features of the channels 24 in the manner described above allows the optimum heat transfer between the blade 10 and the cooling medium to be achieved at specific points of the blade 10 which tend to suffer from extreme thermal stresses and creep and the like.

The width of the ribs 30 along the entire length of the channels 24 is constant at around 1 mm in the disclosed embodiment, but if required the rib width may also be varied.

Reference is now made to Figure 6 in which there is shown a channel 24 of the blade 10 of Figure 1, showing the detail in the region of one of the apertures 28 in the tip 14 of the blade 10. The turbulence promoting ribs 30 are distributed throughout the channel 24 in a manner which ensures that the terminating end of the channel 24 at the tip 14 includes a rib 30 in order to maximise the number of ribs 30 that are present in the channel 24 for a given distribution, and also to maximise the turbulence imparted to the flow of cooling medium.

The preferred method by which the channels 24 and ribs 30 are formed involves a known electrochemical machining process wherein an electrode is used to form the channels 24 and ribs 30 in a two stage operation. The process involves the steps of introducing an electrode to at least one of the blade root 18 and the blade tip 14 of a turbine blade 10 and activating and advancing the electrode to penetrate the blade 10 to form a bore. The ribs 30 are formed by varying the rate of advancement of the electrode such that an increased rate of advancement will produce a smaller inner bore diameter and a reduced rate of advancement will produce a larger inner bore diameter. The electrode is advanced to form a first bore until a predetermined point in the blade 10 is reached when the electrode is withdrawn. The same electrode, or alternatively a different electrode is then introduced to the other of said blade root 18 and blade tip 14 where the process is repeated to form a second bore which merges with the first bore to form a complete channel 24. In a preferred embodiment, the first and second bores are formed to be angularly off-set. In this way, a cooling channel 24 is formed in the blade 10 which includes a plurality of upstanding annular ribs 30 along the length thereof. However, the channel 24 extending through the root portion is smooth and as such the electrode is advanced at a constant rate when forming this particular portion of the channel 24.

The electrochemical process is represented in Figures 7 and 8. Figure 7 is a graphical representation of a plot 50 of the rate of advancement of an electrode 52 (Figure 8) against the displacement of the electrode 52 through the turbine blade 10. Figure 8 is a general representation of the form of a channel 24 created in accordance with the plot 50 of the graph of Figure 7.

In Figure 7 , various steps in the process are represented as I-VI, which are shown as corresponding formed features I-VI in Figure 8. The initial step I involves a relatively large rate of advancement of the electrode 52, which is then reduced in step II to a lower rate of advancement. This results in a rib I defining a first diameter and a recess II defining a larger second diameter (Figure 8) due to the fact that in region II the dwell time of the electrode 52 is greater than in region I, and thus more material may be removed by the electrode 52. The subsequent process can be viewed accordingly and it should be understood from Figures 7 and 8 that a lower rate of advancement of the electrode 52 will result in a larger diameter bore being formed.

It would be apparent to those of skill in the art that the embodiments hereinbefore described are merely exemplary of the present invention and that various modifications may be made thereto without departing from the scope of the invention. For example, the invention is not limited to the specific rib distribution described and shown in Figures 3 to 5 and Table 1; any distribution and rib dimensions and properties may be selected in accordance with factors such as materials and turbine operating temperatures and the like. However, it is a desired feature of the present invention to ensure an improved overall heat transfer between the blade and the cooling medium that the channels extending through the aerofoil section include ribs along the entire lengths thereof.

The turbulence promoting ribs may extend at any angle from the surface of the channels and may include gaps therein such that the ribs define discontinuous annular ribs . Furthermore, the bore may be formed in a single stage operation. Additionally, a plurality of bores may be formed simultaneously by use of a comb of electrodes which are activated and advanced together to form multiple bores .

Claims

1. A turbine blade including an aerofoil section comprising opposing pressure and suction side walls adjoining at leading and trailing edges of the blade, and defining at least one internal cooling channel extending through the aerofoil section from a base to a tip thereof, said at least one channel including a plurality of axially spaced and inwardly projecting ribs located on a wall surface of the channel, said ribs being turbulence promoting ribs and being provided along the entire length of the at least one channel .

2. A turbine blade according to claim 1, further including a root portion, wherein the root portion includes a blade platform, upon which blade platform the aerofoil section is mounted via its base.

3. A turbine blade according to claim 2 , wherein the aerofoil section and root portion are integrally formed.

4. A turbine blade according to claim 2, wherein the aerofoil section and root portion are formed separately and subsequently secured together.

5. A turbine blade according to claim 2 , 3 or , wherein the root portion defines at least one internal cooling channel substantially aligned with the at least one cooling channel extending through the aerofoil section.

6. A turbine blade according to any one of claims 2 to 5, wherein the root portion defines a corresponding number of cooling channels as are defined in the aerofoil section.

7. A turbine blade according to claim 5 or 6, wherein the at least one channel of the root portion is smooth and does not include any turbulence promoting ribs.

8. A turbine blade according to claim 5 or 6, wherein the at least one channel of the root portion includes one or more axially spaced turbulence promoting ribs.

9. A turbine blade according to any one of claims 5 to

8, wherein, in use, a cooling medium may be communicated from the at least one channel of the root portion and into an associated at least one channel of the aerofoil section.

10. A turbine blade according to any one of claims 5 to

9, wherein the at least one channel of the root portion and the at least one channel of the aerofoil section together define an at least one blade cooling channel which extends through the root portion, through the aerofoil section and terminates at the tip of the aerofoil section.

11. A turbine blade according to claim 10, wherein the at least one blade cooling channel is open at the root portion and tip of the aerofoil section, thus forming respective apertures at the root and tip of the blade.

12. A turbine blade according to claim 11, wherein, in use, a cooling medium may be supplied to the blade via the respective aperture in the root portion and exit the blade via the respective aperture in the tip of the aerofoil section.

13. A turbine blade according to claim 10, 11 or 12, wherein the at least one blade cooling channel is substantially aligned with the general longitudinal axis of the turbine blade.

14. A turbine blade according to any one of claims 10 to 13, wherein the at least one blade cooling channel displays a change in orientation at an interface between the root portion and the aerofoil section of the blade.

15. A turbine blade according to any preceding claim, wherein the at least one channel of the aerofoil section has a substantially circular lateral cross-sectional shape.

16. A turbine blade according to any one of claims 5 to 15, wherein the at least one channel of the root portion has the same general lateral cross-sectional shape as the at least one channel of the aerofoil section.

17. A turbine blade according to any preceding claim, wherein the at least one channel of the aerofoil section is of circular cross-section and the turbulence promoting ribs are annular ribs which extend radially inwardly from the wall surface of the channel so as to define annular recesses between adjacent ribs.

18. A turbine blade according to any preceding claim, wherein the turbulence promoting ribs each extend uninterrupted around the inner perimeter of the at least one channel .

19. A turbine blade according to any one of claims 1 to 17, wherein at least one of the ribs includes at least one gap therein.

20. A turbine blade according to any preceding claim, wherein the ribs extend substantially perpendicular from the surface of the at least one channel of the aerofoil section.

21. A turbine blade according to any one of claims 1 to 19, wherein the ribs extend at an acute angle from the surface of the at least one channel of the aerofoil section with respect to the intended flow direction.

22. A turbine blade according to any one of claims 1 to 19, wherein the ribs extend at an obtuse angle from the surface of the at least one channel of the aerofoil section with respect to the intended flow direction.

23. A turbine blade according to any preceding claim, wherein the distribution of the turbulence promoting ribs is such that the terminating end of the at least one channel of the aerofoil section at the tip thereof includes a turbulence promoting rib.

24. A turbine blade according to any preceding claim, wherein the general dimension of the at least one channel of the aerofoil section normal to the intended direction of flow of cooling medium varies along the length thereof .

25. A turbine blade according to any one of claims 5 to 24, wherein the general dimension of the at least one channel of the root portion normal to the intended direction of flow of cooling medium is uniform along the length thereof.

26. A turbine blade according to any preceding claim, wherein the rib pitch varies over the length of the at least one channel of the aerofoil section.

27. A turbine blade according to any preceding claim, wherein the rib height varies over the length of the at least one channel of the aerofoil section.

28. A turbine blade according to any preceding claim, wherein the rib height is greatest and the rib pitch is smallest in a central region of the aerofoil section of the turbine blade.

29. A turbine blade according to any preceding claim, wherein the rib height is smallest and the rib pitch is greatest in cooler regions which require a lower rate of heat transfer.

30. A turbine blade according to any preceding claim, wherein the aerofoil section of the turbine blade defines a plurality of internal cooling channels distributed in a general chord-wise direction of the blade between the leading and trailing edges of the aerofoil section.

31. A turbine blade according to claim 30, wherein the root portion of the blade defines a corresponding number of internal cooling channels arranged to be aligned with a respective cooling channel of the aerofoil section.

32. A turbine blade according to any preceding claim, wherein eight cooling channels are provided.

33. A turbine blade according to any preceding claim, wherein the width of the turbulence promoting ribs is substantially constant over the entire length of the at least one channel of the aerofoil section.

34. A turbine blade according to any preceding claim, wherein the turbine blade is a second stage rotor blade for use in a gas turbine engine.

35. A gas turbine engine including a plurality of turbine blades, at least one turbine blade including an aerofoil section comprising opposing pressure and suction side walls adjoining at leading and trailing edges of the blade, and defining at least one internal cooling channel extending through the aerofoil section from a base to a tip thereof, said at least one channel including a plurality of axially spaced and inwardly projecting ribs located on a wall surface of the channel, said ribs being turbulence promoting ribs and being provided along the entire length of the at least one channel .

36. Electrical generating means including a gas turbine engine, said gas turbine engine including a plurality of turbine blades, at least one turbine blade including an aerofoil section comprising opposing pressure and suction side walls adjoining at leading and trailing edges of the blade, and defining at least one internal cooling channel extending through the aerofoil section from a base to a tip thereof, said at least one channel including a plurality of axially spaced and inwardly projecting ribs located on a wall surface of the channel, said ribs being turbulence promoting ribs and being provided along the entire length of the at least one channel .

37. A method of forming at least one cooling channel in a turbine blade by an electrochemical machining operation, said method comprising the steps of:

(a) providing a turbine blade having an aerofoil section mounted on a root portion;

(b) introducing an electrode to at least one of the root portion and a tip of the aerofoil section;

(c) activating and advancing said electrode to penetrate the blade to form a first bore; and

(d) varying the rate of advancement, of the electrode such that an increased rate of advancement will produce a smaller inner bore diameter and a reduced rate of advancement will produce a larger inner bore diameter, such that a plurality of upstanding ribs are formed between adjacent recesses in the first bore; such that at least one cooling channel is formed in the blade which includes a plurality of upstanding annular ribs along the entire length of said aerofoil section.

38. The method of claim 37 wherein the electrode is advanced in the manner described in step (d) until the electrode emerges from the other of said root portion and tip of the aerofoil section.

39. The method of claim 37, further comprising the steps of:

(e) introducing an electrode to the other of said root portion and tip of the aerofoil section;

(f) activating and advancing the electrode to penetrate the blade to form a second bore; and (g) advancing the electrode until the second bore merges with the first bore.

40. The method of any one of claims 37, 38 or 39, wherein the method is adapted to form a smooth bore in the rib portion of the blade.

41. A method of manufacturing a turbine blade, said method comprising the steps of: forming a blade having an aerofoil section mounted on a root portion; forming at least one cooling channel in said blade by an electrochemical machining process wherein;

(c) the rate of advancement of the electrode is varied such that an increased rate of advancement will produce a smaller inner bore diameter and a reduced rate of advancement will produce a larger inner bore diameter, such that a plurality of upstanding ribs are formed between adjacent recesses in the first bore; such that at least one cooling channel is formed in the blade which includes a plurality of upstanding annular ribs along the entire length of said aerofoil section.

42. The method of claim 41, wherein the electrode is advanced in the manner described in step (c) above until the electrode emerges from the other of said root portion and tip of the aerofoil section.

43. The method of claim 41, further comprising the steps of:

(d) introducing an electrode to the other of said root position and tip of the aerofoil section; (e) activating and advancing the electrode to penetrate the blade to form a second bore; and

(f) advancing the electrode until the second bore merges with the first bore.

44. The method of claims 41, 42 or 43, wherein the variation of the rate of advancement of the electrode is such that the width of the recesses varies along the length of the at least one channel.

45. The method of any one of claims 41 to 44, wherein the variation of the rate of advancement of the electrode is such that the subsequent height of each rib varies along the length of the at least one channel.

46. The method of any one of claims 41 to 45, wherein the rate of advancement of the electrode in the region of the tip of the aerofoil section is such that an upstanding rib is formed at the terminating end of the at least one channel at the tip.

47. The method of any one of claims 37 to 46, wherein multiple bores are created simultaneously using a comb of electrodes in which multiple electrodes are arranged to be activated and advanced at the same time to penetrate the blade and form a plurality of bores.