Turbine vane cooler, turbine vane with said cooler, turbine and aircraft engine with said cooler

TECHNICAL FIELD

The present invention relates to the field of cooling of the turbine blades of the gas turbine engine, in particular an aircraft engine.

More particularly, the invention relates to the field of turbine blade cooling air jets cooling, washing the inner wall of the blade through the jacket positioned within the blade.

In particular, the invention relates to an apparatus of cooling fins disposed on the inner wall of the blades and the openings for supplying jets of cooling air, formed through the shirt.

Furthermore, the invention relates to the relative arrangement of ribs and feed holes in order to improve the cooling of the blade.

The invention also relates to a blade provided with such a cooling device, the blade can be fixed vanes of the distributor or be part of a rolling wheel.

The invention also relates to a turbine comprising at least one cooling device and / or shovel.

Finally, the invention relates to an engine of the aircraft, equipped with at least one above cooling device and / or one above the shoulder blade and / or turbine.

BACKGROUND ART

Known cooling unit blade turbine stage of a gas turbine engine.

The blades are cooled by forced air convection cooling, circulating through the cavity created in the blades. The cooling air is taken from the coldest part of the gas turbine engine, such as a compressor. This air cooling takes place in each blade of the extreme parts of it, such as its outermost radial portion. This cooling air circulates in the blade and extends through the opposite extreme zone, for example via its inner radial edge portion. Sometimes cooling air passes into the blade on its both edge portions.

When the vanes are provided with a jacket disposed in the cavity corresponding blade, cooling air can be circulated as part of the inner jacket to the outer side thereof, through the supply port extending through the jacket.

This cooling air washes the inner wall of the vane cavity through openings in the jacket feed stream in the form of jets and is incident on the wall in different places of contact.

Known technical solution consists in equipping the internal wall of the cavity blade cooling fins to increase the area of ​​heat exchange between the inner wall and flushed with air (US 5,533,864). The presence of these ribs can increase the area of ​​heat exchange between the cooling air jets and the internal wall of the blade.

In cooling devices, which are described above, according to the prior art, the quality of cooling vanes is provided a thermal exchange area between the inner blade wall and the cooling air jets which wash the inside wall, i.e. the number and size of ribs disposed on the inner wall of each blade.

In these known devices, the cooling area of ​​impact of the jets of cooling air with respect to the installation of ribs are arranged randomly. It follows that the cooling of the respective vanes is not uniform in the given zone. In continuation of this, the operations, the most intense in terms of thermomechanical behavior of components, especially in the critical operating point of the turbine engine, the random nature of the location of the ribs with respect to the zones of influence of air cooling can lead to premature wear of the distributor or rolling the turbine wheel.

DISCLOSURE OF INVENTION

The present invention provides an improvement of cooling devices fixed or moving blades of the turbine of the turbojet engine by means of cooling air jets, when each blade has at least a cavity and at least a jacket disposed within the cavity. Cooling is produced by jets of air cooling, washing the inner wall of the blade through the feeding holes formed in the jacket. Cooling fins are arranged on the inner wall of the blade in the region opposite the exit of the jets of cooling air in those places where the jets of cooling air with the wall face.

According to the invention, a special configuration of the cooling fins, positioning with respect to the feeding openings formed in the jacket, so as to overcome the aforementioned drawbacks of known cooling devices.

According to a first embodiment of an apparatus of cooling a turbine blade of a gas turbine engine, the blade is provided with at least a cavity surrounded by an inner wall comprising cooling fins, separated by intervals of pitch p and having a thickness e, and the blade is provided with a jacket located inside the above-mentioned cavity and is traversed by the feed holes with a diameter d, characterized in that when placing the shirts in the cavity vane and critical operating point of the turbine engine, each supply opening shirt is opposite places on the inner wall of the blade, which is located between cooling fins.

Preferably, the shirt was placed in the cavity of the scapula, and in the critical operating point turbomachine pitch p and the thickness e of cooling fins conform against

(1) (p / e) ≥3.

Preferably also, the shirt was placed in the cavity of the scapula, and in the critical operating point turbomachine at least one place on the inner wall of the blade, which is located between the cooling fins is located at least opposite the openings in the jacket.

In other words, when the jacket is in the blade cavity, and critical operating point of the turbine engine, none of the edges of the inner wall is not located opposite the blade openings in the jacket.

According to a second embodiment of the invention relates to a turbine blade of a gas turbine engine includes a cooling device according to the first aspect of the invention.

According to a third embodiment of the invention relates to a turbine of a gas turbine engine comprising at least one blade according to the second embodiment.

According to a fourth embodiment of the invention relates to an aircraft engine, comprising at least one turbine blade according to the second embodiment.

Due to the cooling device according to the invention is achieved a more uniform cooling of the turbine blades than with the prior art cooling devices. As a consequence, the life of the turbine increases.

With a cooling device according to the invention achieved more efficient cooling of the turbine blades than with known cooling devices.

According to another embodiment of the engine of an aircraft engine of an aircraft according to a fifth embodiment of the invention includes a cooling device of a turbine blade according to the first embodiment is characterized in that it comprises a selection device of the cooling air, which cooling air is supplied to the interior of jacket disposed in the cavity, at least one blade according to the first embodiment, the aforementioned cooling air washes and then an inner wall of the blade through executed in the aforementioned shirt feed holes as cooling flow for which the Reynolds number Re, having as characteristic length diameter d feed holes corresponds to the ratio

(2) Re <10,000.

According to yet another embodiment of an aircraft engine, according to a fifth embodiment of the invention includes a cooling device of a turbine blade comprises the above blade provided with at least a cavity surrounded by an inner wall comprising cooling ribs separated from one another at intervals with a pitch p and a thickness e, wherein the aforementioned blade is provided with a jacket, placed inside the aforementioned cavity, and a cross-feed openings having a diameter d, said aircraft engine further includes a selection device of the cooling air, which cooling air flows into the inside of the jacket, the aforementioned cooling air washes then the inner wall of the vane through the feed holes formed in said jacket, and is served as a cooling flow for which the Reynolds number Re is as characteristic length of the diameter d of the transmission holes, and characterized in that, when the shirt is located in the cavity vane and critical operating point of the turbine engine, each supply opening is situated opposite the jacket on the inner wall of the blade, which is disposed between the cooling ribs and the pitch p, the thickness e and Reynolds number Re fit relationship

(1) (p / e) ≥3 and (2) Re <10,000.

It is noted that with the device according to the invention the cooling of the turbine blade cooling modes optimized at the most dangerous with regard to the provisions of details thermomechanical behavior.

Advantageously, the critical operating point turbomachine average coefficient of heat exchange between the flow of cooling fins and cooling was increased by 10% relative to the coefficient of heat exchange that takes place when using a prior art cooling devices.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is further explained by description of nonlimiting embodiments thereof with reference to the accompanying drawings, in which:

Figure 1 represents, in longitudinal section, blade jacketed distributor situated inside the cavity and with supply openings created in the jacket;

Figure 2 presents the same spreader vane, in a cross sectional view along line II-II, indicated in Figure 1;

Figure 3 - schematic diagram illustrating, in longitudinal sectional view, the relative location of the edges of the inner wall and blade holes generated in the jacket;

Figure 4 - is an isometric view of the inner wall edge of the blade.

DETAILED DESCRIPTION OF EMBODIMENTS

As is known, the turbine consists of a sequence of steps, each of which contains a valve and a movable wheel, the distributor is an array of vanes which straighten the airflow over the path of the turbomachine, and the movable wheel carrying the movable vane.

Referring to Figures 1 and 2, consider the turbine blade 10 of the distributor of the turbomachine.

Shoulder 10 is provided with a cavity 12 and an inner wall 14 surrounding the cavity 12. The blade 10 includes a first end 16 and second end 18, which in the example illustrated in Figure 1, its ends are radially outer and radially inward. Both ends 16, 18 are open and communicate with the cavity 12 of the vane 10.

Paddle 10 includes attack edge 20 and a trailing edge 22.

On its inner wall 14 of the blade 10 is provided with cooling ribs 24 which are arranged precisely on the edge 20. Since the attack is illustrated in Figure 4, cooling fins 24 are shown in a detail, cut from the plate and having a rounded profile.

The cooling fins 24 are characterized by the following dimensions:

- The thickness e (see FIG. 4), and

- A step p, which corresponds to the average distance between two adjacent wings cooling.

The cavity 12 is a shirt 26, which is in the form of a prefabricated and welded sheet having a conical shape. The ends of the jacket are located on an inner wall of the jacket 14. The body 26 located between its ends, is not in contact with the inner wall 14. It is by the attack edge 20 of the blade 10 is a gap 40 between the jacket 26 and the inner wall 14. Only several lugs 27 on the square jacket 26 is in contact with the inner wall 14 and perform positioning and supporting cavity 26 in the jacket 12.

The shirt 26 is provided with feed holes 28 which are formed by drilling the wall of jacket 26. These feed holes 28 have a diameter d.

In the example illustrated in Figures 1 and 2, blade 10 also includes brackets 30 which extend through the cavity 14 transversely thereto and the through-holes 32 are located at the trailing edge 22. The brackets 30 and apertures 32 play no role in the present invention.

It is known that during functioning of a gas turbine engine turbine blade 10 of the distributor are cooled by cooling air, drawn in the cooler portion of the turbine engine, such as a compressor, with the introduction of cooling air into the cavity 12 of each blade 10.

In Figure 1 the arrow 100 indicates the flow of cooling air introduced into the jacket 26 disposed in the cavity 12 through first end 16 and the arrow 200 shows part of the flow of cooling air which exits the jacket 26 through the second end 18. This air stream which intersects the jacket 26 allows to supply cooling air of other engine parts.

A substantial part of the flow of cooling air enters the jacket 26 from the first end 16, and it extends from the inner wall 14 of the blade 10 through the feed holes 28 drilled in the jacket 26, for creating transverse flows. These transverse currents are shown in Figure 3 by the arrows 400.

The cooling air exits through the supply port 28 in the form of air jets which wash the inner wall 14 in areas 50 collisions with it.

When the jacket 26 located in cavity 12 and in the functioning of the most rigorous in terms of thermomechanical turbomachine cooling fins 24 and supply port 28 are formed respectively on the inner wall 14 of the blade 10 and the jacket 26 located in the blade 10, thus that the point of impact of the jets 50 of cooling air is disposed between the cooling fins 24, as shown in Figure 3. In other words, the critical operating point of the cooling air stream does not fall on the cooling fins 24, and fall on the area between them. It may happen that only one jet of air cooling or more jets of cooling air will fall at a location between the cooling fins 24. The functioning conditions, the most severe in terms of thermomechanical turbomachine, particularly at the critical point, correspond to high temperatures, lies in the range between 700 ° C and 1100 ° C, wherein the mechanical characteristics of the materials deteriorate. In these most critical operating conditions of the cooling blades 10 should be the most efficient.

It has been observed that the relative positioning of the cooling fins on the inner wall 14 of the blade 10 and feed holes 28 extending through the jacket 26, is optimum when the following relationships are satisfied:

(1) (p / e) ≥3 and

(2) Re <10,000

where p is the average distance between two adjacent cooling fins 24, e represents the average thickness of the ribs 24, d represents the diameter of the feed holes 28, and Re is the Reynolds number in the flow of cooling air at the outlet of the feed holes 28.

The above relates to a turbine blade of a gas turbine engine distributor. Meanwhile, the invention is also applicable to a movable blade of the turbine wheel gas turbine engine.

Application Example

The blade is made of superalloy. The shirt is made of superalloy.

For the flow of cooling air at the feed holes formed in the jacket having a Reynolds number Re = 5000, a critical operating mode corresponding point must have the following dimensions:

p = 3.5 mm

f = 0.5 mm

d = 0.53 mm.

In practice, a digital simulation model is implemented as follows:

- Suggest a "hot" model, that is the proper mode of critical operating points, introducing as parameters to the desired temperature determined depending on the materials used, and selecting the above-mentioned values ​​p, e, d;

- Then go from "hot" model to a "cold", changing the parameters of temperature,

- Then reduced values ​​p, e, d of this "cold" model that corresponds with and relative positions of edges and the supply opening parts manufactured at low temperatures.

Numerical simulations performed on the software-aided design CAO (Conception Assistée par Ordinateur), possessing a module for interfacing with the terms of the temperature. The above example was carried out with the module Scale Factor software-aided design CAO Catia.