CN116104587A - Gas turbine guide of aeroengine - Google Patents

Abstract

The invention relates to the technical field of turbine engines, and discloses a gas turbine guide for an aeroengine; it includes an annular inner ring, an annular outer ring, and guide blades connecting the annular inner ring and the annular outer ring; the annular inner ring is provided with a first air film hole , the annular outer ring is provided with a second air film hole, the guide vane has a hollow cavity connecting the inner surface of the annular inner ring and the outer surface of the annular outer ring, and the buckle connection in the hollow cavity acts as vibration damping and airflow cooling for the guide vane The impingement tube, the front end of the guide is socketed with the flame tube, the end of the guide is connected with the inner casing of the turbine, and the annular inner ring, annular outer ring and guide vanes are integrally formed. Through the above design, the thermal stress of the guide is smaller, the structure is more stable, and the service life is longer. The ring-shaped inner ring is provided with a vibration-damping guide groove, and a variable resistance plate with its vibration-damping function is installed, which can not only reduce noise, but also improve Vibration damping further increases guide life.

Description

Gas turbine guide for aero-engine

technical field

The invention relates to the technical field of turbine engines, in particular to a gas turbine guide for aeroengines.

Background technique

Existing aeronautical turbine engine gas turbine guides are widely used in precision casting of high-temperature alloys and film cooling, and the gas deflects at an accelerated rate in the guide channel, whizzing past, and has a strong aerodynamic excitation force. The gas turbine guide of a large engine is assembled into a whole ring by fan-shaped blocks, and the blade of a single guide can be oriented crystallized or single crystal, and each guide is thermally deformed and vibrates, and rubs against each other, so that the vibration stress is significantly reduced. Crack growth, but the number of parts is large, the assembly relationship is complicated, and the manufacturing cost is high. The gas turbine guide for small and medium-sized aviation turbine engines adopts integral welding or integral precision casting to reduce the number of parts. However, the overall precision cast gas turbine guide has high rigidity, high natural vibration frequency, small vibration damping, and large thermal stress, and its cracking tendency is obviously higher than that of the integrally welded gas turbine guide.

The prior art CN201420831817.9 discloses a turbine guide, including an inner ring (10), an outer ring (20) and several guide blades (30), and several mounting holes ( 21), the mounting holes (21) are evenly distributed around the axis of the outer ring (20), one end of the guide vane (30) is fixedly connected to the outer circumferential surface of the inner ring (10), and the guide vane (30) The other end is inserted into the installation hole (21), and the number of the guide blades (30) is the same as the number of the installation holes (21).

Among the turbine guides disclosed in the prior art, only a basic mechanism of the turbine guide is disclosed, and no technical solution for reducing thermal stress is disclosed, that is, there are still problems in the prior art that the guide is easy to crack and be scrapped due to large thermal stress. technical flaws.

Contents of the invention

The technical problem solved by the invention is to overcome the defects in the prior art and provide a gas turbine guide for an aero-engine with small thermal stress and long service life, so as to solve the technical problem of large thermal stress of the guide.

The object of the present invention is achieved through the following technical solutions:

Disclosed is a gas turbine guide for an aero-engine, the guide includes an annular inner ring, an annular outer ring, and guide blades connecting the annular inner ring and the annular outer ring; the annular inner ring is provided with a first air film hole , the annular outer ring is provided with a second air film hole, and the guide vane is provided with a third air film hole; the guide vane has a hollow cavity connecting the inner surface of the annular inner ring and the outer surface of the annular outer ring, The hollow cavity is buckled with an impact tube that acts as vibration damping and airflow cooling for the guide blades. The front end of the guide is socketed with the flame tube, and the end of the guide is connected with the inner casing of the turbine. The annular inner ring, the annular outer ring and the guide blades are integrally formed.

Preferably, the number of guide vanes is 18.

Preferably, the end of the annular inner ring adjacent to the inner casing of the turbine is provided with a first lace, and the first lace is provided with fork ears adapted to the flame tube.

Preferably, a second lace is provided on the end of the annular outer ring adjacent to the turbine inner casing, and a pin hole mounting seat adapted to the turbine inner casing is arranged on the second lace.

Preferably, the inner surface of the hollow cavity is provided with a first protrusion, and the guide vane abuts against the outer surface of the impact tube through the first protrusion; the outer surface of the impact tube is provided with a second Two protrusions, the second protrusion is fixedly connected to an edge of one end of the hollow cavity.

Preferably, the outer surface of the impingement tube is in close contact with the inner surface of the hollow cavity.

Preferably, the impact tube is formed by bending an iron-based superalloy plate around, and the bent impact tube has a slit.

Preferably, the impact tube is provided with a fourth air film hole.

Preferably, the annular inner ring is provided with a guide groove for installing the varistor, one end of the guide groove is connected to the end of the annular inner ring close to the inner casing of the turbine, and the other end is connected to the end of the annular inner ring away from the turbine. The ends of the inner receiver are not connected.

Preferably, the varistor includes a long bar installed in the guide groove and a horizontal bar connected to the long bar, and the horizontal bar is connected to the inner case of the turbine.

Preferably, the varistor is made of an elastic varistor.

Compared with the prior art, in the guider of the present invention, the front end of the guider is socketed with the flame tube, and the end of the guider is connected with the inner casing of the turbine; the annular inner ring, the annular outer ring and the guide vanes constitute its main structure , and the three are integrally formed, through this design, the number of connected parts that are directly used for connecting and fixing parts is reduced, the overall structure is more refined, and the assembly time is also reduced. At the same time, cold air blows to the inner wall from the small hole at the rear stop of the inner flame tube, and a first air film hole is set in the annular inner ring to form a cold air film boundary layer in the inner flow channel; the cold air enters the flow channel from the gap in the outer wall stop , the annular outer ring is provided with a second air film hole to form a cold air film boundary layer of the inner flow channel; and the guide vane is designed as a hollow structure with a hollow cavity, and a pair of Impingement tube with guide vanes for vibration damping and airflow cooling. Through the above design, the thermal stress of the whole guide is reduced.

It has the following beneficial effects:

1) The annular inner ring, annular outer ring and guide vanes are integrally formed, which reduces the number of engine drawings and parts, reduces complicated welding procedures, facilitates mass production, and reduces manufacturing costs.

2) Higher gas temperature and gas pressure are allowed to improve the efficiency of the heat engine.

3) The reliability is improved, the failure rate is reduced, the maintenance and repair work is reduced, and the service life of the engine overhaul is extended.

4) The thermal stress is lower, the thermal cycle residual deformation is smaller, and the coaxiality is better.

5) The vibration damping is larger, the vibration stress is smaller, the service life of the guide vane is longer, and the life of the guide is longer.

Description of drawings

Fig. 1 is the structural representation of the gas turbine guider of a kind of aero-engine of the present invention;

Fig. 2 is the sectional view of the gas turbine guider of a kind of aero-engine of the present invention;

Fig. 3 is a schematic structural view of an impingement tube in a gas turbine guide of an aero-engine of the present invention;

Fig. 4 is the structural schematic diagram of the varistor in the gas turbine guider of a kind of aero-engine of the present invention;

Fig. 5 is a schematic diagram of the connection between the gas turbine guider and the turbine inner casing of an aero-engine according to the present invention.

Detailed ways

In order to clearly illustrate the technical features of this solution, the present invention will be described in detail below through specific implementation modes and in conjunction with the accompanying drawings.

In the following description, many specific details are set forth in order to fully understand the application, but the application can also be implemented in other ways different from those described here, therefore, the protection scope of the application is not limited by the specific details disclosed below. EXAMPLE LIMITATIONS.

In addition, in the description of the present application, it should be understood that the terms "center", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal" ", "Top", "Bottom", "Inner", "Outer", "Axial", "Radial", "Circumferential" and other indicated orientations or positional relationships are based on the orientations or positional relationships shown in the drawings , is only for the convenience of describing the present application and simplifying the description, but does not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and thus should not be construed as limiting the present application. In addition, the terms "first" and "second" are used for descriptive purposes only, and cannot be interpreted as indicating or implying relative importance or implicitly specifying the quantity of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of these features. In the description of the present application, "plurality" means two or more, unless otherwise specifically defined.

In this application, terms such as "installation", "connection", "connection" and "fixation" should be interpreted in a broad sense, for example, it can be a fixed connection or a detachable connection, unless otherwise clearly specified and limited. , or integrated; it can be a mechanical connection, an electrical connection, or a communication; it can be a direct connection or an indirect connection through an intermediary, it can be the internal communication of two components or the interaction relationship between two components . Those of ordinary skill in the art can understand the specific meanings of the above terms in this application according to specific situations.

In the present application, unless otherwise clearly specified and limited, a first feature being "on" or "under" a second feature may mean that the first and second features are in direct contact, or that the first and second features are indirect through an intermediary. touch. In the description of this specification, descriptions referring to the terms "one embodiment", "some embodiments", "example", "specific examples", or "some examples" mean that specific features described in connection with the embodiment or example , structure, material or characteristic is included in at least one embodiment or example of the present application. In this specification, the schematic representations of the above terms are not necessarily directed to the same embodiment or example. Furthermore, the described specific features, structures, materials or characteristics may be combined in any suitable manner in any one or more embodiments or examples.

Example 1

As shown in Figure 1, a gas turbine guide for an aero-engine is disclosed, and the guide includes an annular inner ring 1, an annular outer ring 2, and guide vanes 3 connecting the annular inner ring 1 and the annular outer ring 2; the annular inner ring 1 is provided with There is a first air film hole 4, a second air film hole 5 is provided on the annular outer ring 2, and a third air film hole is provided on the guide vane 3; the guide vane 3 has an inner surface connecting the annular inner ring 1 and the annular outer ring 2 The hollow cavity 6 on the outer surface is snap-connected with an impact tube 7 that acts as vibration damping and airflow cooling on the guide vane 6. The front end of the guide is socketed with the flame tube, and the end of the guide is connected with the turbine inner The casing 13 is connected, and the annular inner ring 1, the annular outer ring 2 and the guide vanes 3 are integrally formed.

In this embodiment, the front end of the guide is socketed with the flame cylinder, and the end of the guide is connected with the inner casing 13 of the turbine. The guide mainly includes an integrally formed annular inner ring 1 , an annular outer ring 2 and guide vanes 3 , and the guide vanes 3 are connected between the annular inner ring 1 and the annular outer ring 2 . In this embodiment, specifically 18 guide vanes may be provided, and of course the number of guide vanes may be more than 18. Specifically, it can be processed by means of precision casting, and its advantage is that the favorable performance effect of the overall precision casting gas turbine guide is to allow higher gas temperature (about 1450K) and gas pressure to improve the efficiency of the heat engine. At the same time, the beneficial effect of the overall precision casting gas turbine guide is to reduce the engine drawing number and the number of parts, reduce the complicated welding process, facilitate mass production, and reduce manufacturing costs. Appropriate guide vane wall thickness and strength adapt to higher loads, making the turbine do more work, and only one stage of gas turbine is enough to drive the compressor; and the beneficial effect of the integral precision casting gas turbine guide is the improvement of reliability, Reduce the failure rate, reduce maintenance and repair work, and extend the life of the engine overhaul; and the integral precision casting gas turbine guide is convenient to use the alloy powder laser rapid printing molding process to manufacture the guide test piece for testing the engine. Through precision casting, not only the production efficiency of the guide is improved, but also the connection is more stable, which avoids cracks during the use of the guide. The annular outer ring 2 and the annular inner ring 1 are trumpet-shaped, the annular inner ring 1 is provided with a first air film hole 4 , and the annular outer ring 2 is provided with a second air film hole 5 . The cold air is blown from the small hole in the rear seam of the inner flame tube to the first air film hole 4 surrounding the front end of the annular inner ring 1 to form the boundary layer of the cold air film in the inner flow channel; The hole enters the flow channel and blows to a circle of second air film holes 5 arranged around the front end of the annular outer ring 2 to form a cold air film boundary layer of the outer flow channel. The guide vane is also provided with a third air film hole, and the guide vane 3 has a hollow structure with a hollow cavity 6, and an impact tube 7 is arranged in the hollow cavity 6 to increase vibration damping and airflow cooling. Through this design, the thermal stress of the guide blade is reduced, the thermal stress is lower, the residual deformation of the thermal cycle is smaller, and the coaxiality is better. Therefore, the structural stability of the guide vane during operation is ensured, and the service life of the guide is finally improved.

Example 2

As shown in Figure 1, a gas turbine guide for an aero-engine is disclosed, and the guide includes an annular inner ring 1, an annular outer ring 2, and guide vanes 3 connecting the annular inner ring 1 and the annular outer ring 2; the annular inner ring 1 is provided with There is a first air film hole 4, a second air film hole 5 is arranged on the annular outer ring 2, and a third air film hole is arranged on the guide vane; The surface of the hollow cavity 6, the hollow cavity 6 is buckled and connected with the impact tube 7 for vibration damping and airflow cooling of the guide vane 6, the front end of the guide is socketed with the flame tube, and the end of the guide is connected with the turbine inner machine The cassettes 13 are connected, and the annular inner ring 1, the annular outer ring 2 and the guide vanes 3 are integrally formed.

The difference between this embodiment and Embodiment 1 is that: the end of the annular inner ring 1 adjacent to the inner casing 13 of the turbine is provided with a first lace 8, and the first lace 8 is provided with a fork ear 9 adapted to the flame tube; A second lace 10 is provided on the end of the annular outer ring 2 adjacent to the turbine inner casing 13 , and a pin hole mounting seat 11 adapted to the turbine inner casing 13 is arranged on the second lace 10 . When the guide sleeve is respectively socketed with the flame tube and the inner casing 13 of the turbine. The overall precision-cast gas turbine guide bears the same aerodynamic torque as the gas turbine, so the force on the outer ring is relatively large, and there are 16 pins and screws connected to bear the shear stress. Considering the strength at high temperature, the radial fork lugs are arranged on the turbine inner casing 13 with a slightly lower temperature, and the second lace 10 is punched with precision round holes for mounting pins. Screws and self-locking nuts gently tighten the guide and the turbine inner casing 13, and the turbine inner casing 13 remains motionless when the guide expands with heat and contracts with cold. Through the above design, the connection between the guide sleeve and the flame tube and the inner casing 13 of the turbine is more stable.

The overall precision casting gas turbine guide is made of nickel-based high-temperature alloy material. According to the alloy melt flow performance, the casting generally has a thickness of 1.6-1.5mm, the thinnest blade body wall thickness is 1.2-1.0mm, and the thickest installation fork lug, round The thickness of the ring edge is 4-3mm, and the final thickness of the machined surface is 3mm. Good steel is used on the cutting edge. The nickel-based high-temperature alloy overall precision-cast gas turbine guide only includes guide vanes 3, annular outer ring 2, annular inner ring 1, fork ears 9 and lace. The flow channel wall is designed to be as short as possible. The first Lace 8 and the second flower 10 are also arranged with multiple lace gaps. The streamlined design makes the overall precision casting small in size and light in weight, which is very beneficial to the entire casting process of the nickel-based superalloy from smelting to cooling, and the structure of the final guide is simpler.

Example 3

As shown in Figure 1, a gas turbine guide for an aero-engine is disclosed, and the guide includes an annular inner ring 1, an annular outer ring 2, and guide vanes 3 connecting the annular inner ring 1 and the annular outer ring 2; the annular inner ring 1 is provided with There is a first air film hole 4, a second air film hole 5 is arranged on the annular outer ring 2, and a third air film hole is arranged on the guide vane; The surface of the hollow cavity 6, the hollow cavity 6 is buckled and connected with the impact tube 7 for vibration damping and airflow cooling of the guide vane 6, the front end of the guide is socketed with the flame tube, and the end of the guide is connected with the turbine inner machine The cassettes 13 are connected, and the annular inner ring 1, the annular outer ring 2 and the guide vanes 3 are integrally formed.

The difference between this embodiment and Embodiment 1 is that: the inner surface of the hollow cavity 6 is provided with a first protrusion, and the guide vane 3 abuts against the outer surface of the impingement tube 7 through the first protrusion. The outer surface of the impact tube is provided with a second protrusion. When the impact tube 7 is inserted into the hollow cavity 6, the second protrusion is clamped and connected to one end edge of the hollow cavity, and the first protrusion on the hollow cavity 6 will abut against the outer surface of the impingement tube 7 , so that the guide vane 3 is fixedly connected with the impingement tube 7 . In order to insert the impingement tube 7 into the hollow interior 6, the first protrusion can be provided as a round first protrusion. The outer surface of the impact tube 7 is in close contact with the inner surface of the hollow cavity 6, thereby increasing the contact area between the two, thereby increasing the vibration damping of the guide vane 3. Specifically, the thickness of the impact tube 7 and the guide vane 3 can also be adjusted. The thicknesses of the relative hollow cavities 6 are set to be the same. In order to further protect the guide vane 3, aluminized chromium and aluminum silicon can also be arranged on the outer surface of the guide vane 3. The impact tube 7 can be bent from an iron-based superalloy plate with high temperature resistance and elasticity. The bent impact tube has slits and has a certain degree of elasticity. At the same time, the fourth air film hole 70 can also be set on the impact tube 7, so as to accelerate the heat transfer of the guide vane, improve the service life of the guide vane 3, and finally increase the service life of the guider.

On the surface of the hollow cavity 6 of the guide vane 3, there are many hemispherical convex points in contact with the impact tube 7, which also play the role of vibration damping and airflow cooling. Greater vibration damping, less vibration stress and longer guide life. The percussion sound of the gas turbine guider in the assembled state is dull and rapidly attenuated, so that it can also play a role in noise reduction.

Example 4

As shown in Figure 1, a gas turbine guide for an aero-engine is disclosed, and the guide includes an annular inner ring 1, an annular outer ring 2, and guide vanes 3 connecting the annular inner ring 1 and the annular outer ring 2; the annular inner ring 1 is provided with There is a first air film hole 4, a second air film hole 5 is arranged on the annular outer ring 2, and a third air film hole is arranged on the guide vane; The surface of the hollow cavity 6, the hollow cavity 6 is buckled and connected with the impact tube 7 for vibration damping and airflow cooling of the guide vane 6, the front end of the guide is socketed with the flame tube, and the end of the guide is connected with the turbine inner machine The cassettes 13 are connected, and the annular inner ring 1, the annular outer ring 2 and the guide vanes 3 are integrally formed.

The difference between this embodiment and the above-mentioned embodiments is that: the annular inner ring 1 is provided with guide grooves for installing the varistor 12 , and in this embodiment, there are nine uniformly arranged guide grooves. One end of the guide groove is connected to the end of the annular inner ring 1 near the turbine inner casing 13, and the other end is not connected to the end of the annular inner ring 1 away from the turbine inner casing 13. The varistor 12 includes a long bar 120 installed in the guide groove and a horizontal bar 121 connected to the long bar 120. The horizontal bar 121 is connected with the inner casing 13 of the turbine. Specifically, the varistor 12 can be set as an elastic variable blocking film. In comparison, the disadvantages of integral precision casting gas turbine guides are that casting defects are difficult to avoid, large thickness, high rigidity, large thermal stress, high natural vibration frequency, and small self-vibration damping. The knocking sound of the overall precision casting blank is crisp and lingering, which is not a good thing for the hot end parts. Therefore, 9 guide grooves (0.3mm wide) are prefabricated on the annular inner ring 1, which is equivalent to prefabricated cracks, and the thermal strain deforms the width of the narrow slits to prevent the guide vanes 3 from being pulled apart. Cutting the slits reduces the stiffness and the natural frequency. Cutting slits in the inner wall and inner ring will not propagate into cracks. Insert a 0.2mm thick elastic varistor 12 in the guide groove, when the gas turbine guide vibrates. The elastic varistor 12 rubs in the guide groove to attenuate the vibration amplitude, thereby reducing the dynamic stress and prolonging the cycle life of the gas turbine guide.

Apparently, the above-mentioned embodiments are only examples for clearly illustrating the technical solution of the present invention, rather than limiting the implementation of the present invention. For those of ordinary skill in the art, other changes or changes in different forms can be made on the basis of the above description. It is not necessary and impossible to exhaustively list all the implementation manners here. All modifications, equivalent replacements and improvements made within the spirit and principles of the present invention shall be included within the protection scope of the claims of the present invention.

Claims (10)

1. A gas turbine guide for an aero-engine, characterized in that the guide includes an annular inner ring, an annular outer ring and guide vanes connecting the annular inner ring and the annular outer ring; the annular inner ring is provided with The first air film hole, the second air film hole is provided on the annular outer ring, the third air film hole is provided on the guide vane; the guide vane has an inner surface connecting the annular inner ring and an outer surface of the annular outer ring The hollow cavity is provided with an impact tube for vibration damping and airflow cooling of the guide vane, the front end of the guide is socketed with the flame tube, and the end of the guide is connected with the inner casing of the turbine , the annular inner ring, the annular outer ring and the guide vanes are integrally formed. 2. A gas turbine guide for an aero-engine according to claim 1, wherein the number of said guide vanes is 18. 3. A gas turbine guide for an aero-engine according to claim 1, wherein a first lace is arranged on the end of the annular inner ring adjacent to the inner casing of the turbine, and the first lace is provided with There are prong lugs to fit the flame barrel. 4. A gas turbine guide for an aero-engine according to claim 1, wherein a second lace is arranged on the end of the annular outer ring adjacent to the inner casing of the turbine, and a second lace is arranged on the second lace. There are pin-and-hole mounts that fit the turbine inner case. 5. A gas turbine guide for an aero-engine according to claim 1, wherein a first protrusion is provided on the inner surface of the hollow cavity, and the guide vane passes through the top of the first protrusion. against the outer surface of the impact tube; the outer surface of the impact tube is provided with a second protrusion, and the second protrusion is fastened to an edge of an end of the hollow cavity. 6 . The gas turbine guide for an aero-engine according to claim 1 , wherein the outer surface of the impingement tube is in close contact with the inner surface of the hollow cavity. 7 . 7 . The gas turbine guide for an aero-engine according to claim 1 , wherein the impingement tube is formed by bending an iron-based superalloy plate, and the impingement tube after bending has a slit. 8 . The gas turbine guide for an aero-engine according to claim 1 , wherein a fourth air film hole is arranged on the impingement tube. 9. A gas turbine guide for an aero-engine according to claim 1, wherein a guide groove for installing a varistor plate is provided on the annular inner ring, and one end of the guide groove is connected to the annular inner ring. The end of the ring is close to the inner casing of the turbine, and the other end is not connected to the end of the annular inner ring away from the inner casing of the turbine. 10. A gas turbine guide for an aero-engine according to claim 9, wherein the varistor comprises a long bar installed in the guide groove and a horizontal bar connecting the long bar, so The horizontal bar is connected with the inner casing of the turbine, and the varistor is elastic.

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