Air cooled turbine vanes

(54) Name: Device at a cooled turbine joint vane

7513487-4

The present invention relates to a device for a cooled turbine (inlet) articulated vane, comprising a hollow, stream-shaped vane blade having an inlet or front edge, having a plurality of cooling air holes in gas connection with the cavity portion of the leading edge, and having a pressure side and a vacuum or suction side, a plurality of cooling air holes in gas communication with the leading edge cavity portion, a cooling air splitter insert with one side located closest to the pressure side and its other side located closest to the suction side; when the insert is positioned in the blades cavity, and a device located in the blades cavity, which includes a sealing strip or rib located on the suction side of the paddle blade between its cooling air holes and the front air cooling holes, and a second sealing strip or rib which is the pressure side of the vane blade in the vicinity of the front air cooling hole, so that the cooling air flow to the front air cooling hole is kept separate from the cooling air flow to the suction side cooling air hole; whereby the insert closest to the suction side respectively. pressurized sides are provided for sealing abutment against the first sealing strip or sealing strip, respectively. the second sealing strip.

A cooled turbine joint vane of this type is previously known, for example, from U5 patent 3,799,696, preferably Figures 4 and 5.

In this turbine guide vane, the suction and pressure sides of the streamlined vane blade are cooled by jet cooling, while the leading edge of the vane blade is cooled by film or boundary layer cooling.

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Film or boundary layer cooling requires a carefully determined but relatively small pressure drop through the cooling air holes on the front edge of the vane. If the pressure drop is too large, the released air flow will penetrate the passing means and is diverted in the flow direction along with the combustion gases, thus, without forming any cooling boundary layer on the vane surface. On the other hand, if the pressure drop is too small, the hot combustion gases will penetrate the cooling air layer and cause harmful heating of the paddle material. Radiant cooling, on the other hand, requires a large pressure drop through the cooling air holes in the suction and pressure sides to accelerate the air flows to high impact velocities against the paddle wall. In order to establish the required pressure drop, the pressure inside the cavity must be considerably higher than the pressure of the working medium against which the cooling air jets are emitted.

In order to obtain an effective combination of the two cooling concepts, it is important to seal sealingly the front air cooling holes from the cooling air holes of the pressure and suction sides. For this purpose, according to the above-mentioned publication, a tubular insert is used which is in communication with the first and second sealing strips in the cavity of the vane blade. It has been found that this tubular insert does not effectively prevent flow communication between the cooling air holes in the leading edge and the cooling air holes in the pressure and suction sides, so that the object of the present invention is to improve the known turbine joint vane and obtain better sealing between jet cooling and film cooling air flow.

The invention is characterized in that the cooling air splitter insert consists of a substantially U-shaped, flexible part, the legs of which are forced into sealing abutment against the sealing strips in functionally dependent on the internal pressure in the blade cavity.

This flexible, U-shaped insert is forced against the sealing strip in the cavity of the paddle blade by compressive forces in the cavity during engine operation, thereby effectively separating the jet cooling and film cooling outflows from each other. The U-shaped flexible insert provides the desired substantial improvement over the tubular insert according to US Patent 3,799,696. The tubular insert is inserted into the cavity of the paddle blade with a press fit, and since the insert itself is rigid, no line contact is obtained between the sealing strips and the pipe and the pipe by the compressive forces during operation cannot be forced into line contact with the sealing strips. According to the invention, this sealing line contact is obtained between sealing strips and U-shaped insert, whereby the cooling air flow along the wall of the paddle blade cavity between strips and insert is prevented.

An example chosen, particularly suitable embodiment of the invention will be described in the following with reference to the accompanying drawings, in which:

Fig. 1 is a simplified axial section through a portion of a gas turbine engine and particularly illustrates a guide vane in the inlet of the turbine;

Fig. 2 is on a larger scale a cross-section through said guide vane, taken along

7513487-4 line 2-2 in Fig. 1;

Fig. 3 is a similar cross-section through the same articulated vane and illustrates its internal device under pressure, and Fig. 4 is a perspective view of the vane of Fig. 2 with parts broken away.

Part of the turbine section of a gas turbine engine 10 is shown in axial section in Fig. 1.

An inlet guide vane 12 and a turbine vane 14 are located in an annular through-flow duct 16 for combustion gases coming from a combustion chamber 18. The inlet led vane 12 is included in a circular wreath of such vanes located in the same position in the axial flow in the gas inlet. Similarly, the turbine blade is included in a turbine blade pulley located in the through-flow channel immediately downstream of the pulley blade pulley. Each articulated paddle has a radially outer foot 20 and a radially inner foot 22, which carry between them the streamlined paddle blade 24, which thus spans the distance between the two paddle feet.

As can be seen in Fig. 2, each articulated blade has a leading edge chamber 26 and a trailing edge chamber 28. The leading edge 30 of the blade is facing upstream and is provided with front edge cooling sheath 32 which is distributed therein between the inner and outer feet of the blade. The front air cooler hole connects the front edge chamber 26 to the annular flow passage 16. The vane blade further has a hook edge 34 which has a rear edge air hole 36. This cooling air hole is included in a row of such holes, distributed along the trailing edge between the inner and outer paddle feet. The trailing holes connect the trailing chamber 28 to the annular flow passage 16. Each guide vane blade has a pressure side 38 which shows a first series of cooling air holes 40 connecting the leading edge chamber 26 with the annular flow passage, and a second series of cooling air holes 42 connecting the rear edge chamber 28 with this throughflow channel for the working gas stream. Each streamline-shaped paddle blade further exhibits a suppressor¹ or suction side 44, which has a first series of cooling air holes 46 connecting the leading edge chamber 26 to the flow passage 16, and a second series of cooling air holes 48 connecting the rear edge chamber 28 to this channel 16.

The leading edge chamber 26 has a pressure wall 50, which has a sealing strip 52 and a support bar 54 projecting inwardly from said wall 50. In Figure 2, only one carrier bar 54 arranged on the pressure wall is shown, but in reality a series of such bosses are distributed in the same in the axial position along the chamber wall 50. The leading edge chamber 26 further has on the vane or suction side of the vane a wall 56, which has a sealing strip 58 and a support bar 60. In fig. 2 Although only one such support bar 60 is shown, but in reality a series of such barbs are distributed in the same axial position along the chamber wall 56. The trailing edge chamber 28 has on the printing side of the vane a wall 62, which has a sealing strip 64 and a carrier bar projecting inwardly from this wall 66. several such bosses are distributed in the same axial position along this wall 62, although only one boss 66 is visible

7513487-4 in Fig. 2. The trailing edge chamber also hooks on the suction side of the vane a wall 68 which is provided with a sealing strip 70 and a support bar projecting from the wall inwardly. A plurality of such bosses are distributed in the same axial position along this wall 68, although only a cam 72 is visible in Fig. 2. The leading edge and trailing edge chambers are separated from each other by a beam 74 provided with a plurality of respectively. chamber insertion pushers 76. A leading edge insert 78 and a trailing edge insert 80, which have approximately U-shaped cross-sectional profile, are located in the leading edge chamber 26, respectively. the trailing edge chamber 28. Each of the insert's one leg 82 is located opposite and respectively. the wall of the chamber closest to the pressure side, while the second leg 84 of the inserts is located opposite and respectively. the wall of the chamber closest to the suction or vacuum side. A plurality of beam cooling holes 86 are provided in the front and rear edge inserts, 78 and 80.

When the gas turbine engine is in operation, air is compressed in its compressor portion and supplied to the combustion chamber 18, in which a portion of the pressurized gases is mixed with fuel to a combustible mixture which is combusted to increase the gas energy's moving energy. The combustible mixture should be burned at high temperatures to reduce the quantity of off-burning hydrocarbons leaving the combustion chamber together with the combustion gases. The desired combustion temperatures are far higher than the maximum permissible temperature to which longer downstream metallic parts may be exposed, so that dilution air from the engine's compressor portion is fed into the combustion chamber's outlet area. In a typical modern engine, the dilution air is mixed with the combustion gases to reduce the maximum temperature of local gases flowing into the turbine under the engine take-off state of the aircraft lifting to about 1650 ° C at a static pressure of 23.4 kp / cm 2. The inlet joint paddle ring, which is in the annular flow through of the combustion gases to the turbine inlet, directs the gas stream at an optimal angle into the turbine paddle ring 14. Each of the paddle sheave's linear-shaped blades 24 is specially formatted to direct the flow of gas into the turbine inlet. A concave surface on the pressure side 38 of the scabbard blade is hit by the backward flowing combustion gases and provides them with a peripherally directed flow component. A convex surface on the oppression or suction side 44 of the nearest paddle vane is located opposite the pressure side of the streamlined blade of the former paddle sheave, so that a paddle channel acting as part of a turbine inlet is formed therebetween. In the aforementioned typical engine, the absolute static pressure of the gases flowing along the pressure side of the vane blade is 23.2 kp / crn. The absolute static pressure on the opposite suction side of the adjacent paddle blade is 19 kp / cm 2 and the absolute static pressure at the trailing edge of the paddle blade is 17.7 kp / crn.

The requirement for cooling of the articulated blade is most critical in the area of ​​the leading edge 30, where the working temperature and pressure are highest. Cooling air, which is removed

7513487-4 from the compressor of the gas turbine engine described above, the leading edge chamber 26 is supplied at a pressure of about 23.6 kp / cm, or 99% of the working medium pressure at the leading edge of the articulated blade during high power development, e.g. when starting and relieving. Frontal layer cooling at the leading edge is well known as an extremely effective means of avoiding overheating of the metal parts in this area. When boundary layer cooling is applied, a continuous stream of cooling air with a low flow velocity flow is released through the cooling air holes 32 at the leading edge of the blade. The cooling air thus released is deflected by the hot working media gases in the reverse direction along the blade surfaces to be cooled. If the pressure drop through the cooling air holes 32 in the leading edge of the blade is too large, the cooling air flow rate becomes too large, whereby the cooling air will enter the working media stream during turbulence, mixing with its hot gases and thereby destroying the cooling capacity of the boundary layers. On the other hand, if the flow rate is too low, the working media gases will displace the air boundary layer and come into contact with the metal vane blade metal surfaces.

Once the correct ratio of cooling air flow to the working media flow and the correct flow rates has been determined using methods well known in the art, devices of the present invention are used to provide and maintain these determined flow conditions.

The front edge insert 78, which has a substantially U-shaped profile, bends the front edge holes 32 and the first series of cooling air holes 40 on the pressure side of the vane. Although the first series of cooling air holes on the vane side of the vane is lacking in some vane designs, however, they are shown in this particularly advantageous embodiment for increasing the boundary layer of cooling air along the vane blade pressure. · Side, where temperatures are highest. The pressure side cooling air holes are separated from the air holes in the leading edge to take advantage of the controlled flow through the leading edge holes through the device constructed in accordance with the present invention.

It is also well known to those skilled in the art that jet cooling of the inner wall surfaces of the vane blade is an effective complement. to the film or boundary layer . Contrary to the pressure differential requirements for boundary layer cooling, the jet cooling systems require a relatively large pressure drop between the cooling air chambers and the surface to be cooled, so that the cooling air can be accelerated to a velocity at which the air jets hit the surface to be cooled. At the same time, the radiant cooling air must be delivered at a relatively low pressure in order to maintain the relatively large pressure drop between the respective airflows. air comb axle and cooled surface. In the present particularly advantageous embodiment, a region of considerably reduced pressure is located in the working passage through the workpiece adjacent a vacuum or suction wall surface of the guide vane blade, and the radiant cooling air is consequently discharged at this location.

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Radiation cooling of the inner wall surfaces and boundary layer cooling of the outer wall surfaces are effectively combined in arrangement according to the particularly advantageous embodiment of the invention. Interfacial cooling air emitted through the leading edge holes is kept separate from the radiant cooling air which flows through the first series of cooling air holes 46 arranged in the vacuum side into the annular flow passage 16. It is imperative that a considerable stream of cooling air from the leading edge chamber flows out through the leading edge air holes, so that a barrier of boundary layer cooling air is formed around the leading edge of the vane. If the cooling air flow here is interrupted locally, the vane blade is exposed to the hot working gases and overheated until finally destroyed. By separating the supply of air to the leading edge, it is ensured that the correct proportion of cooling air is directed to the cooling air hole of the leading edge rather than to the area with lower pressure along the suction or suction side of the vane blade. Thus, any lack of cooling air to the leading edge area is avoided.

The shank 82 of the leading edge insert 78 closest to the pressure side and the shank 84 closest to the suction side shaft are deflected in the leading edge chamber towards the pressure side sealing strip 52 and 52, respectively. against the suction side sealing strip 58. As can be seen in Fig. 2, the sealing strips are located on either side of the front air cooling hole 32, and in the present embodiment also outside the first series of the pressure side cooling air holes 40. The series of abutment arches are arranged along the inner wall surfaces of the vane blade to keep the deflected insert at a predetermined distance from corresponding chamber walls in the vane blade. Air flows at great speed across the space thus delimited and hits and cools the inner wall surfaces of the blade. As the chamber is relieved from pressure, the legs of the U-shaped insert spring back to an internal position at a distance from the carrier connectors. The seal between the insert legs and corresponding sealing strips also exists in the pressure relieved state, and this seal does not interfere with any premature contact of the insert legs with the support connectors when the chamber is pressurized. In addition, the carriers keep the U-shaped insert at a distance from the inner wall surfaces of the paddle blade, so that a series of passages are formed in between. These passages direct the cooling air stream along the inner wall surfaces to convectively cool the paddle walls. The leading edge insert is made of thin metal sheet, preferably in the thickness range of 0.20 to 0.25 mm, but a thickness of 0.10 to 0.50 mm may be useful, depending on the pressure differences, temperature and length of the insert legs. The insert itself is elastically flexible and made free of moldings, flanges or other surface roughness, which would give this sheet metal increased shape stiffness. The insert is dimensioned so that press fit is obtained between the sheet insert and the sealing strips against which the insert rests, and the insert is subjected to a clamping pressure only between contact points of

7513487-4 sealing strips when the engine is not working. The U-shaped insert represents a significant technical advance over the prior art known in the art.

The trailing edge insert 80 operates in the trailing edge chamber 28 in a similar manner to the leading edge insert of the leading edge chamber. However, in this particularly advantageous embodiment, the pressure drop between the inside and outside of the walls of the rear cooling air chamber is greater than the pressure drop between the inside and outside of the walls of the front cooling air chamber, so it has been found that the plate thickness of the rear insert should be 0.27 to 0. , 33 mm.