Cooled turbine guide blade and turbine having said blades

FIELD: gas turbine engines.

SUBSTANCE: guide blade of gas turbine engine comprises front and rear edges, pocket, wing back, as well as opened perforated insert, inlet orifice to supply cooling air inside insert and outlet orifice to remove cooling air portion from blade. The insert defines annular cavity arranged between outer side wall thereof and inner blade wall. Blade has bridges located in cavity part between the insert and inner surface of rear edge to increase blade rigidity. One insert end is fixed to blade. Another one is installed so that the insert end may slide along inner edge of blade under the action of mutual thermal insert expansion relatively inner blade wall. Orifices are created in opened insert so that the orifices are located only in two insert areas. Orifices of the first group are opposite to inner surface of front edge. Orifices of the second group are opposite to inner surface of rear edge to prevent air jet impingement upon bridges.

EFFECT: increased cooling ability.

5 cl, 6 dwg

The technical field to which the invention relates.

The present invention relates to turbine blades of gas turbine engine, in particular of the guide vanes has a built-in cooling circuit.

The level of technology

Known gas turbine engines include a combustion chamber in which fuel is mixed with air, after which the resulting mixture is burned. The gases resulting from this combustion, flow in the direction of the rear end of the combustion chamber, and then do a high-pressure turbine and low-pressure turbine. Each turbine contains one or more rows of stationary guides) blades (form to allow the device), interspersed with one or more rows of moving (working) blades (installed on the rotary disks), all of these blades distributed circumferentially around the turbine rotor. The turbine blades are exposed to high temperature combustion gases reaching levels far exceeding limit values, which vanes are in direct contact with the gases can withstand, without damage, which leads to a reduction of the lifetime.

A known method of solving this problem is to provide such a blade internal cooling circuit, designed to lower the temperature of the blades by the organization aimed circulation of cooling air within the blade. In the walls of the blades is provided with holes designed to create on the surface of the blades protective air layer.

Figure 4 and 5 shows the guide vane of known construction, which is cooled by means of an open liner, which is currently used in the nozzle apparatus of some aircraft engines.

The blade 10, containing a hollow pen 12, located between the external (banding) shelf 14 and the inner shelf 16 (shelf lock), also contains liner 18, which limits the peripheral annular cavity 20 located between the inner wall of the pen and the external lateral surface of the liner. The outer end portion 18A of the liner tightly attached to the outer flanges of the blades by welding or soldering, and its inner portion 18B is inserted into the guide zone 16A of the inner shelf blades with a certain clearance required for installation of the liner and its sliding displacement under the influence of thermal expansion. Spacers 22 provided on the inner wall or formed by protrusions in the liner, provide a constant distance between the liner and the inner wall and lintel 23 attach the blade stiffness in the remaining part of the cavity 20.

Outdoor liner 18 is equipped with multiple holes. Thus, the cooling air entering the t source of air under pressure, as a rule, from the gas turbine compressor, enters the outer shelf 14 through the inlet 24. Next, the air enters the liner 18, and part of it goes through multiple openings of the liner, forming a peripheral cavity 20 air jets. These jets cool the inner wall of the pen 12 in clashes with her face with the ridges 23 and output via a calibrated outlet openings 26 provided in the rear edge or in a trough (internal concave surface) of the pen, forming a protective air layer along the entire length of this rear edge. The remaining part of the cooling air is discharged through the inner shelf 16, passing through it and cooling it, and then leaving the blade through the hole 28 and doing other engine components, such as disk drives, require cooling.

This known construction is generally satisfactory. However, the limitation of the pressure of the air holes of the liner and, indirectly, the cross-section of the leakage reduces the efficiency of cooling. In addition, the walls of the vanes, in particular, near its front edge there are significant local variations of the temperature gradient, leading to the emergence of radial stresses. These stresses are in extreme operating modes have a negative impact on the work of lo is ADI.

Disclosure of inventions

Thus, the problem to which the present invention is directed, is to eliminate the disadvantages associated with significant variations of the temperature gradient through the creation of the guide vanes of a gas turbine engine, cooled by means of an open liner and characterized by significantly reduced radial stresses acting on the blade. The invention also includes a turbine of a gas turbine engine equipped with such guide vanes cooled by means of an open liner.

In accordance with the invention the solution of this problem is achieved by the creation of the guide vanes of a gas turbine engine having a front edge, a back edge, a trough and the back of the pen, as well as outdoor perforated liner, which limits the annular cavity located between the outer wall of this liner and the inner wall of blades, and jumpers, located in part of the cavity between the liner and the inner surface of the rear edges to give the blade stiffness. There is also an inlet for supplying cooling air to the inside of the liner and the outlet for the output of the cooling air from the blade. At one end of the liner is rigidly connected to the blade, and the other egomanic can slide on the inner edge of the scapula due to relative thermal expansion of the liner and the inner wall of the blade. The blade according to the invention is characterized by the fact that in the open liner made holes only in two specific areas of the liner. Holes of the first group are located opposite the inner surface of the leading edge and the holes of the second group are located opposite the inner surface of the rear edge. The orifices of the second group are arranged so as to prevent collision of the air jets with the ridges of the scapula.

Thus, increasing the cooling front and rear edges of the blades and limiting the impact of the jets of air jumpers, achieve a significant reduction in local temperature gradients and reduce the level of axial mechanical stresses in the blade.

Ideally, each of the first and second groups of holes contains no more than three rows of holes of the liner, and in a typical embodiment, one row of holes of the insert.

The second group of holes are preferably located in the area of the liner corresponding to the air jets coming out of these holes, the shortest path between the liner and the inner surface of the rear edge. In addition, in the proposed configuration, the second group of holes may be located in an area of the liner corresponding to the air jets coming out of these holes, the shortest path is between the liner and grooved outlet holes, made in the trough of the pen to form a protective air layer along the entire length of the rear edge.

Brief description of drawings

Other properties and advantages of the present invention will become apparent from the following descriptions with reference to the accompanying drawings, illustrating an example embodiment of the invention, without introducing any limitations. In the drawings:

- figure 1 represents a longitudinal section of the guide vane of the turbine according to the invention,

on figa presented in an enlarged scale part 1,

- figure 2 represents a transverse section of the blade of figure 1,

- figure 3 shows a graph of the temperature change of the outer part of the blade of figure 2,

- figure 4 and 5 depict two orthogonal cuts known guide vane of the turbine.

Information confirming the possibility of carrying out the invention

Figures 1 and 2 show a cooled blade 10, which is, for example, the guide vane of the turbine of the gas turbine engine of the present invention. The blade is formed with a hollow pen 12 established between the external (banding) shelf 14 and the inner shelf 16 (shelf lock)attached to the housing (not shown) of the turbine by means of the external shelf, which forms the outer wall of the channel of flow of combustion gases through the turbine. The inner wall of the channel current of gases on risovana inner shelf of this blade.

Relative to the direction of flow of gases, indicated in figure 1 by the arrow, usually define the front edge 12A, the rear edge 12B of the trough (inner concave side) 12C and back (outer convex side) 12D of this pen.

Such guide vane is exposed to combustion gases having a very high temperature, and therefore requires cooling. For this purpose, the blade 10 according to the invention, known as blades, contains at least one open perforated liner 18, which in one of its radial ends of the supplied cooling air and which restricts peripheral annular cavity 20 located between the inner wall of the vanes and the outer side wall of the liner. The outer radial portion 18A of the liner tightly attached to the outer shelf 14 by welding or soldering, and its inner radial portion 18B is inserted into the guide zone 16A of the inner shelf 16 with a certain annular gap. This clearance is required for installation of the liner and its sliding displacement in the process, taking into account the uneven temperature changes of the various components of the scapula and, consequently, their relative extension, and education zone leakage of the cooling air. The cooling air after passing through the liner is shown, in addition, calibrated through the haunted outlet 26, made in the trough of the pen to form a protective air layer along the entire length of the rear edge 12B. In the external and internal shelves are provided, respectively, the inlet opening 24 for air and an outlet opening 28 for output air to circulate cooling air.

In accordance with the invention serves to focus the holes in the insert is in two areas that are particularly sensitive to the effects of the hot combustion gases moving in the channel current of gases, namely, in the zones of the liner opposite front and rear edges of the pen blades.

As shown in figure 2, the front edge 12A is cooled by jets of air coming from the first group of holes 30 of the liner opposite the leading edge, thus providing cooling in this zone of the inner surface of the pen airflow. Ideally, this first group of holes contains a single row of holes.

The rear edge 12B is cooled by air by means of a calibrated outlet openings 26, in which the air flows from the second group of holes 32 of the liner located essentially opposite the inner surface 12 Bi trailing edge and, thus, providing cooling in this zone of the inner wall of the pen airflow. Ideally, this second group of holes also will gain one row of holes. In this case, as shown in an enlarged scale on figa, these holes are located so that, for jets of air coming out of these holes, provided the shortest path between them and grooved outlet holes 26 and prevented the collision of the jets of air from the inner wall of the scapula and with the ridges 23 of the blade. Thus, the air passing through the calibrated outlet remains relatively cold, because only part of it, formed by the jets passing through the first group of holes 30, is heated by contact with the hot inner wall near the front edge, and with the trough and the back of the pen. At the same time, another part is formed by the jets, coming right through the second group of holes 32 with the heating due to the absence of collisions with the inner wall and with the ridges before it reaches the calibrated outlet.

The results of the implementation of the invention from the point of view of change of temperature is presented in figure 3 corresponding to the radial scan of the cross section depicted in figure 1 and 2 vanes of the guide vane low pressure. The section plane passes through the middle part of the channel current of gases, i.e. in the zone of maximum temperature combustion gases.

The first curve 34 corresponds to the hypothetical to the hence, adaptation, in which the number of holes of the liner is extremely small and, therefore, virtually all of the cooling air exits through the leakage area between the liner and the inner shelf (i.e. equivalent to the cross section of the holes of the liner is much less than the cross section of the leak). It may be noted that in this limit configuration adequate cooling wall is provided only near the trailing edge (sq), and the wall near the front edge (PC) practically is not cooled. As a result, the temperature in this zone is very close to the temperature of combustion gases and, therefore, is within the acceptable for this material range.

The second curve 36 corresponds to a known configuration in which the openings of the liner distributed over the entire surface of the liner. In this configuration, the leading edge is adequately cooled, and its temperature is below the limit of destruction pen blades. However, there are significant fluctuations in temperature, creating a large temperature gradients that give rise to local mechanical stress, reducing the strength of the blades and, consequently, reduce its service life.

In contrast, using the present invention (the third, the solid curve 38) these oscillations are significantly reduced, and the temperature of the leading edge and trailing edge is reduced by the number of the degrees. Typical values of this reduction is from 4 to 8 degrees. In addition, on both sides from the front edge, at a distance not exceeding half of the distance between the front and rear edges (i.e. for values of the abscissa <0,5), can be achieved very significant improvement of the temperature gradient of the scapula. More precisely, for values of the abscissa in the range from approximately 0.2 to 0.4 improvement Δ temperature gradient for pen trough, reaching 60 degrees.

Thus, this configuration of the holes of the insert part, excessively cooled using known technologies, heated stronger, and part of the cooled enough, additionally cooled. This achieves the best balance with respect to the temperature gradients. Although the description were, essentially, only the guide blade turbine gas turbine engine, it is evident that the design of the guide vanes are cooled by means of an open liner that can easily be used in the stator of the compressor of the gas turbine engine, as well as the blades of the casing of the gas turbine engine.

1. Guide vane of a gas turbine engine having a front edge (12A), trailing edge (12V), trough (12C) and back (12D) of the pen, as well as an outdoor pen is arrowny liner (18), which limits the annular cavity (20)located between the outer side wall and the inner wall of blades (10), jumpers (23), located in part of the cavity between the liner and the inner surface (12B_i) rear edges to give the blade stiffness, the inlet opening (24) for supplying cooling air to the inside of the liner and the outlet (28) for the output of the cooling air from the blade, and one end (18A) of the liner is rigidly connected to the blade, and the other end (18V) is slidable on the inner edge (16A) of the scapula due to relative thermal expansion of the liner and the inner wall of the blade, characterized in that in the open liner made holes in only two areas of the liner, and the holes (30) of the first group are located opposite the inner surface (12A_i) the leading edge and the holes (32) of the second group are located opposite the inner surface of the rear edge in such a way as to prevent collision of the jets of air bridges.

2. The blade according to claim 1, wherein each of the first and second groups of holes of the liner contains no more than three rows of holes (30A, 30B, 30C) of the insert.

3. The blade according to claim 2, wherein each of the first and second groups of holes of the insert contains one row from which Erste liner.

4. The blade according to claim 2 or 3, characterized in that the second group of holes is located in the area of the liner, suitable for air jets coming out of these holes, the shortest path between the liner and the inner surface of the rear edge.

5. The blade according to claim 2 or 3, characterized in that the second group of holes is located in the area of the liner, suitable for air jets coming out of these holes, the shortest path between the pad and grooved outlet holes (26), made in the trough of the pen to form a protective air layer along the entire length of the rear edge.

6. Turbine gas turbine engine, characterized in that it contains guide vanes are cooled by means of an open liner made according to any one of claims 1 to 5.