Fluid cooled vane assembly

2.071.665

19) FRENCH REPUBLIC

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(Use only for filing and reproduction orders.)

NATIONAL INSTITUTE OF INDUSTRIAL PROPERTY

National registration number

70.31836

PARIS (To be used for annual payments, requests for official copies and all other correspondence with ΙΊ.Ν.Ρ.Ι.) © PATENT D'INVENTION

FIRST AND UNIQUE PUBLICATION

22) Date of filing ................................

Date of grant decision ......

Publication of the issue ..............

September 1, 1970, 3:30 p.m.

August 1971.

BOPI - "Lists" n. 37 of 17-9-1971.

(δΐ) International classification (Int. Cl.) .. F 02 c 7/00 // F 01 d 5/00.

(7 ^ Applicant: Company known as: GENERAL ELECTRIC COMPANY, residing in the United States of America.

(73) Holder: idem (7l) (77) Agent: Cabinet Beau de Loménie, Ingénieurs-Conseils, 55, rue d'Amsterdam, Paris (8). (54) Dawn for gas turbines and device for cooling it.

(72) Invention of:

(33) (32) (3d) Conventional priority: Patent application filed in the United States of America on December 1, 1989, n. 881.254 named after Robert John Smuland, Ned Alexander Hope and James Edgar Sidenstick.

Sale of booklets at IMPRIMERIE NATIONALE, 27. rue de la Convention - PARIS (15^e)

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The present invention relates to gas turbine engines, and relates more particularly to an improved blade intended for such engines, as well as a device for cooling it.

It is well known that the efficiency of a gas turbine engine is related to the operating temperature of the turbine and that the efficiency of the engine can be increased, theoretically.₃ by the increase, of this temperature. However, in practice, the maximum operating temperature of the turbine is generally limited by the resistance of the various elements of the turbine to high temperatures, among which the blades or fins of the turbine generally limit the maximum temperature most.

We already know different types of hollow blades receiving relatively cold air inside which is discharged or sucked from the oppressor and which are intended to increase the maximum operating temperature of the turbine and thus allow an increase, theoretically possible. , the efficiency of the turbine. As the use of pressurized air from the compressor represents a load or in itself constitutes a reduction in the efficiency of the word air, it is important, in such devices, that the heat exchange properties of the blade be such as little refrigerant as possible.

Another problem encountered with such blades is the tendency for the side walls to balloon and crack under the effect of the pressure difference between the inside and the outside of the blade.

The object of the invention is to produce a hollow blade of low weight, having better resistance to bloating and having better properties for heat exchange and cooling.

Another object of the invention is to provide a device for cooling a turbine which is designed for the efficient use of a refrigerant to keep the operating temperatures of various elements of the turbine at low values.

According to the invention, these results are obtained by means of a blade of the type which can be assembled with identical blades in the form of a crown, and which comprises a hollow aerodynamic part extending approximately radially through a current.

31836 hot gas between parts forming bases spaced from each other. The side walls of the hollow aerodynamic part carry reinforcing ribs arranged in cords and projecting inside the aerodynamic part. High heat exchange rates between a refrigerant - as a rule derived from the engine compressor "and the aerodynamic part are obtained, in part, by the provision of at least one inner tubu ™ lining which is intended to receive and directing the refrigerant in the form of a large number of jets with high flow velocity against the side walls of the aerodynamic part. The latter preferably comprises a spacer extending between the side walls inside the aerodynamic part, this spacer dividing the interior into a chamber adjacent to the leading edge, hereinafter called the front chamber and a chamber adjacent to the edge of leak, hereinafter called rear chamber, each comprising a tubular interior lining.

The blade or the adjacent support structure of the engine carries elements directing the refrigerant against the parts forming the bases of each blade so as to ensure high rates of heat exchange between the refrigerant and these bases. The latter have orifices through which refrigerant is directed towards the rear edge of the bases and into the stream of hot gas for additional cooling of the vane bases and for the formation of a film of refrigerant. for cooling elements of the turbine which are mounted downstream.

Other characteristics and advantages of the invention will emerge more clearly from the description which follows, given solely by way of nonlimiting example, as well as from the appended drawings in which:

- Figure 1 is a perspective view, partly in section, of an embodiment of a blade according to the invention;

- Figure 2 is a cross-sectional view along line 2-2 of Figure 1 j

FIG. 3 is a cross-sectional view of part of the turbine of a gas turbine engine comprising a blade like that shown in FIG. 1 as well as the cooling device according to the invention:

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- Figure 4 is a cross-sectional view of a portion of the turbine of a gas turbine engine, showing another embodiment of a blade and a cooling device according to the invention.

Similar parts of the preferred embodiments described below are designated by the same references.

FIG. 1 represents an air-cooled hollow blade of the type intended to be assembled with identical blades in the form of a crown. The blade is designated by 10 and comprises a part 12 forming an internal base and a part 14 forming an external base which are connected together by at least one aerodynamic part 16 which extends approximately radially. The aerodynamic part 16 has a leading edge 18 which is spaced from a trailing edge 20 and which is connected to the latter by a concave side wall 22 and a convex side wall 24. A spacer or partition 25 extends between the side walls 22, 24 &amp; the interior of the aerodynamic part 16, and this partition divides the interior into a front chamber 26 and a rear chamber 28. Tubular linings 30 and 32, having side walls 34 roughly corresponding to the shape of the side walls 22 and 24 of the aerodynamic part, see Figure 2, are appropriately fixed in the front and rear chambers, so that their walls 34 are located a short distance from the side walls of the aerodynamic part. Each interior trim has a large number of orifices 36 directing a refrigerant - for example a fluid derived or sucked from the compressor of the gas turbine engine against the leading edge 18 and the side walls 22 and 24 in the form of a large number of jets at high flow speed, so as to ensure a high heat transfer rate between the refrigerant and said parts of the blade.

Devices, in the form of an internal perforated baffle 40 and an external perforated baffle 42, ensure effective cooling of the internal base 12 and the external base 14, because they direct the coolant against said manifolds in the form of a large number of jets with high flow speed, so as to ensure a high rate of heat exchange between the refrigerant and the heaters.

As is best seen in Figures 1 and 3, a passage 37

31836 passes through the interior base 12 and the interior baffle 40 for the supply of the refrigerant to the lining 30 of the front chamber, and a passage 38 passes through the exterior base 14 and the external baffle 42 for the supply of the refrigerant 'to the lining 30 of the front chamber and the lining 32 of the rear chamber.

The lower perforated baffle 40, see FIGS. 1 and 3, is disposed between flanges 44 and 46, directed radially towards the inside of the base 40, and it is fixed to these flanges by welding, soldering or by other means appropriate. Similarly, the perforated baffle 42 is disposed between the flanges 48 and 50 which form part of the outer base 14 and which are directed radially outward. The baffle 42 is fixed on these edges in the same way as the baffle 40 on the edges 44 and 46.

The rear edge 50 of the external base 14 has a large number of orifices 52, intended to direct the rebounded refrigerant "sanf" of the external base 14 in the direction of the arrows shown in FIG. 3 towards the rear edge of 1 '' external base. The inner base 12 similarly has a large number of orifices 54 which are intended to direct the rebounding fluid from the base 12 towards the rear edge thereof. During operation, the blades 10 direct a working fluid or a stream of hot gas 56 coming from a source such as a combustion chamber 57 towards a ring of turbine blades 66 and cause a predetermined vortex angle C (measured from an axial plane comprising the lines 57 of FIG. 1). In order to limit the mixing and the reduction of the pulse of the hot gas stream 56 to a value as low as possible, since they reduce the efficiency of the turbine, the orifices 52 preferably form an angle with said axial plane, if although the refrigerant leaving these orifices meets the current of hot gas 56 at a predetermined angle A which corresponds approximately to the vortex angle Ç of the hot gas. Similarly, the orifices 54 are of. preferably formed such that the refrigerant joins the hot gas stream 56 at a predetermined angle B which corresponds approximately to the vortex angle C of the hot gas. ''

Although the angles of flow A and B of the refrigerant, determined by the orifices 54 and 52, are preferably equal to the vortex angle C, they can also be less than the latter.

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For example, the Applicant has discovered that, for a vortex angle C of approximately 73 °, a flow angle B of approximately 65 ° and a flow angle A of approximately 55 ° give good results if the 'account is taken both of the cooling efficiency during the transport of the refrigerant through the orifices, of the manufacturing possibilities, and of the subsequent use of the refrigerant leaving the orifices for the formation of a film of refrigerant, as described in more detail below.

As shown in Figure 2, the rear chamber of the aerodynamic part has several longitudinal reinforcing ribs 58 which are spaced from each other, which start from the side walls 22, 24, protrude inward and extend from 1 'spacer or partition 25 to a point 60 located downstream of the interior lining 32 where opposite ribs are connected to each other to thereby form, in cooperation with the partition 25 and the side walls 22, 24, a sort of ribbed box of great rigidity but of low weight, all around the lining 32. The Applicant has found that such a ribbed box structure is particularly effective in reducing the stresses in the side walls 22, 24, therefore also their tendency to bloating and cracking under the effect of the pressure difference between the refrigerant and the stream of hot gas, without such a structure disturbing the cooling by jets of the side walls 22, 24 or reducing the efficiency thereof.

It is possible to provide, in addition, a large number of rods 62 each establishing a connection between the side walls 22 and 24 downstream of the interior lining 32, in order to further strengthen the blade 10 in the area of ​​the rear chamber, and in order to improve the cooling possibilities of the assembly. A large number of orifices 64, intended for the formation of a film of coolant, can be provided, in a manner known per se, in the leading edge 18 and the side walls 22 and 24 in order to ensure a continuous flow of refrigerant through the front chamber 26, and for the formation of a film of refrigerant exiting along the exterior surface of the aerodynamic part 16 in order to further cool the whole of the blade 10.

Ports 91 in the trailing edge similarly ensure the continuous flow of refrigerant through the rear chamber, from

31836 same as the additional cooling of the trailing edge 20.

FIG. 3 represents a part of the turbine of a gas turbine engine in which a ring of blades 10 is provided upstream of a ring of turbine blades 66 which each extend approximately radially from of a rotor 68 of a turbine, up to the proximity of nesting elements 70. It will be readily understood that the trajectory of the stream of hot gas 56 through the turbine shown in FIG. 3 has a roughly annular shape and is defined, in part, by an inner surface 72 of the outer base 14 of the blade , an inner surface 74 of the fitting elements 70, an inner surface 76 of the inner base 12 of the blade and a part 78 forming the base of a fin.

During operation, a suitable refrigerant, for example a fluid drawn in or derived from a compressor of the gas turbine engine, is supplied through suitable passages, for example passages 77 and 79 shown in FIG. 3 and in FIG. 4 to the perforated baffles 40 and 42, Part of the refrigerant arriving through the passage 79 crosses the perforations 80 of the baffle 40 is directed against the external surface of the external base 14 and then flows through the orifices 52 to meet with the stream of hot gas 56.

Another part of the refrigerant is directed into the interior linings 30, 32 through the passage 38. Due to the escape of part of the refrigerant through the orifices 52, where it joins the hot gas stream 56, as indicated schematically by the arrows in FIG. 3, the hot gases of the stream 56 are prevented from enter the space 82 formed between the external base 14 and the adjacent interlocking elements 70 and, moreover, this part of the refrigerant forms a film over the entire internal surface 74 of the interlocking elements 70, thus lowering the temperature thereof.

Similarly, part of the refrigerant leaving the passage 77 flows through the perforations 80, is directed against the internal base 12 and then flows through the orifices 54 to the stream of hot gas 56, thus cooling further the rear part of the base 12, 'and forming a protective film of refrigerant on the base 78 of the fin. Another part of the refrigerant leaving the passage 77 and have time introduced through the passage 37 into the interior lining 30.

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As best illustrated in Figure 3, the refrigerant inside the lining 30 of the front chamber is directed against the leading edge and the side walls 22 and 24 of the aerodynamic part 16 of the blade in order to '' ensure uniform and efficient cooling of these surfaces. Additional cooling is obtained due to the fact that the refrigerant passes through the orifices 64 whose conformation is such that a film of refrigerant is formed on the external surface of the aerodynamic part of the blade. Similarly, the refrigerant inside the lining 32 of the rear chamber is directed against the side walls 22 and 24 so as to produce a high rate of heat exchange between the refrigerant and these walls. The refrigerant rebounding from these surfaces then flows axially towards the rear through the chambers formed between the reinforcing ribs 58, around the rods 62, then escapes through the orifices 91 of the trailing edge and meets the current. hot gas.

FIG. 4 represents another embodiment of a blade and of a cooling device according to the invention. The baffles · perforated · 40 and 42 are placed more towards the outside, with respect to the corresponding bases 12 and 14, and are suitably fixed to the support structure of the engine, for example by the fastening elements 86 and 88. It will also be noted that the lining 30 of the front chamber of the embodiment shown in FIG. 4 receives only refrigerant passing through the baffle 40 through an appropriate passage, similar to passage 37 shown in FIG. 3, while the lining 32 of the rear chamber receives only refrigerant passing through the external baffle 42.

The example shown in FIG. 4 comprises a screen 90 which is mounted outside the baffle 42 and which is intended to retain the entrained particles which could block the orifices 36 or 64, thus improving the reliability of the assembly and the efficiency of the cooling device. During operation, the refrigerant leaving the passage 79 crosses the screen 90 and the perforations 80 of the baffle 42, then is directed against the external surface of the external base 14. As indicated by the arrows shown in FIG. 4, part of the rebounding refrigerant from the base is then directed into the lining 32 of the rear chamber, while another part is directed through the orifices 52 in the

31836 stream of hot gas 56, as described above in relation to the embodiment shown in FIG. 1. Similarly, the refrigerant leaving the passage 77 is directed against the interior surface of the interior base 12 through the baffle 40, and part of the rebounding fluid is then directed into the lining 32 of the rear chamber, while the remaining part passes through the orifices 54 and sweeps the rear edge of the internal base 12 as described previously in relation to the embodiment shown in FIG. 1.

The invention is not limited to the embodiments described and those skilled in the art may make various modifications to it without departing from its scope.

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