Integrated test structure for blade disc of aero-engine

Abstract

The invention relates to the technical field of aviation equipment test, and discloses an aero-engine bladed disc integrated test structure, which comprises a double-radial-plate turbine disc mounted on a rotating shaft; the bottom of the turbine blade is provided with a front air guide hole corresponding to the front edge and the chord middle area of the blade and a rear air guide hole corresponding to the tail edge of the blade; the bottom, corresponding to the turbine blade, of the double-radial-plate turbine disc is provided with an air cavity and divided into a first air cavity and a second air cavity through a baffle, the bottom of the turbine blade is inserted into the air cavity, the first air cavity corresponds to the front air guiding hole, the second air cavity corresponds to the rear air guiding hole, the first air cavity is communicated with cold air, and the second air cavity is communicated with the hollow disc cavity through a through hole. The disc center communicates with cold air. According to the invention, a complex cooling flow path of a real double-radial-plate turbine disc of an aero-engine can be simulated, and a cooling process that cold air passes through the double-radial-plate turbine disc to supply air to blades can be simulated, so that a blade-disc integrated test experiment is realized, and experimental research is carried out on the turbine blades and the double-radial-plate turbine disc at the same time.

Description

Technical Field

[0001] This invention relates to the field of aviation equipment testing technology, and in particular to an integrated test structure for aero-engine bladed disks. Background Technology

[0002] In high-temperature engine components, the double-spoke turbine disk is subjected to the coupled effects of extreme thermal and mechanical loads over long periods. Its thermal management efficiency and structural integrity have a decisive impact on the overall engine lifespan. Therefore, experimental research on double-spoke turbine disks has become an indispensable part of aero-engine development. By constructing a high-fidelity test environment, key data such as the temperature field distribution characteristics of the double-spoke turbine disk under actual operating conditions, the flow heat transfer efficiency of the cooling air system, and the air supply effect of the film cooling holes / slots can be accurately obtained. These experimental results not only verify the accuracy of thermodynamic models and simulation analyses but also provide a basis for structural optimization and cooling scheme design of the double-spoke turbine disk, possessing significant engineering value for overcoming the technical bottlenecks of hot-end components in high thrust-to-weight ratio engines.

[0003] Currently, mainstream double-spoke turbine disk test setups are primarily designed for traditional single-spoke structures. Due to the simplicity of the disk structure, its cooling flow path is also relatively simple. For single-spoke and double-spoke turbine disks, the cooling gas mainly cools and seals the inlet of the double-spoke turbine disk and supplies air to the turbine blades. However, as aero-engines develop towards higher thrust-to-weight ratios, double-spoke turbine disks have gradually become a research hotspot due to their unique structural advantages. While maintaining equivalent structural strength, they significantly improve overall heat exchange efficiency by increasing the heat exchange surface area and optimizing the flow path layout. However, due to the complexity of the cooling flow path in double-spoke disks, an additional cooling flow path to the turbine disk core is added compared to the original single-spoke disk. Existing test setups are insufficient to meet these research needs, and this lack of testing capability severely restricts the engineering application of double-spoke turbine disks.

[0004] Furthermore, current experimental tests mostly separate the testing of the double-spoke turbine disk and the blades. However, in actual engines, the cooling air is supplied to the blades via the double-spoke turbine disk, and the two affect each other. Therefore, it is necessary to conduct integrated blade-disk testing experiments, while simultaneously carrying out experimental research on the turbine blades and the double-spoke turbine disk.

[0005] Therefore, how to provide a test device suitable for double-spoke turbine disks and capable of conducting integrated blade disk testing is a problem that urgently needs to be solved by those skilled in the art. Summary of the Invention

[0006] The purpose of this invention is to provide an integrated test structure for aero-engine bladed disks to solve the problems existing in the prior art.

[0007] To achieve the above objectives, the present invention provides an integrated test structure for aero-engine bladed disks, comprising:

[0008] A double-spoke turbine disk is mounted on a rotating shaft. The double-spoke turbine disk has a disk center and a hollow disk cavity, and the disk center and the hollow disk cavity are connected.

[0009] The turbine blade has a front air intake corresponding to the leading edge and the chordal region of the blade, and a rear air intake corresponding to the trailing edge of the blade. The double-spoke turbine disk has an air chamber at the bottom of the turbine blade, which is divided into a first air chamber and a second air chamber by a baffle. The bottom of the turbine blade is inserted into the air chamber. The first air chamber corresponds to the front air intake, and the second air chamber corresponds to the rear air intake. The first air chamber is connected to the cold air, and the second air chamber is connected to the hollow disk cavity through a through hole. The disk cavity is connected to the cold air.

[0010] Furthermore, it also includes:

[0011] A cover plate is installed at the front end of the double-spoke turbine disk. The cover plate has a first receiving hole, one end of which is connected to the cold air and the other end is connected to the first air chamber.

[0012] Furthermore, the double-spoke turbine disk extends axially along the shaft to form a flange edge, the flange edge is connected to the disk center, and the flange edge is provided with a second receiving hole for communicating with cold air.

[0013] Furthermore, it also includes a turbine guide vane assembly, the turbine guide vane assembly comprising:

[0014] Turbine guide vanes are mounted on the machine body; the machine body is provided with a first pre-swivel nozzle corresponding to the edge of the double-spoke turbine disk, through which cold air is blown to the edge of the double-spoke turbine disk and forms an air seal in front of the double-spoke turbine disk; the machine body is provided with a second pre-swivel nozzle corresponding to the first receiving hole, through which cold air is sent into the first receiving hole; the machine body is provided with a third pre-swivel nozzle corresponding to the second receiving hole, through which cold air is sent into the second receiving hole.

[0015] The casing has an air vent that communicates with cold air, and the casing defines an inlet and outlet chamber. The inlet chamber is connected to the air vent, a first pre-rotation nozzle, a second pre-rotation nozzle, and a third pre-rotation nozzle.

[0016] Furthermore, the housing includes:

[0017] A front bearing housing is used to support the front bearing. The inlet chamber is located in the front bearing housing, and the air vent is located on the front bearing housing.

[0018] A top casing is disposed on the outside of the turbine guide vanes and turbine blades, and the top casing is connected to the front casing;

[0019] A rear load-bearing housing for supporting the rear bearing is connected to the top housing; the shaft is installed between the front bearing and the rear bearing.

[0020] Furthermore, the front bearing and the rear bearing are positioned on the rotating shaft by bearing retaining rings.

[0021] Furthermore, a turbine disk rear baffle is provided at the rear end of the double-spoke turbine disk, and the turbine blades are axially fixed on the double-spoke turbine disk by the cover plate and the turbine disk rear baffle.

[0022] Furthermore, a thermocouple and a Pitot tube are provided in the inlet chamber; multiple thermocouples are arranged on the front surface and the center surface of the double-spoke turbine disk; multiple thermocouples are arranged in the middle of the turbine blades; and a flow meter is arranged in the air vent.

[0023] Furthermore, the rotating shaft is provided with a protrusion and a disc retaining ring. The protrusion is located at the front end of the double-spoke turbine disk, and the disc retaining ring is located at the rear end of the double-spoke turbine disk. The double-spoke turbine disk is positioned on the rotating shaft by the protrusion and the disc retaining ring.

[0024] Furthermore, the protrusions have a toothed structure, allowing cold air to flow from the gaps between adjacent protrusions to the center of the disc.

[0025] The present invention discloses the following technical effects:

[0026] This invention can simulate the complex cooling flow path of a real double-spoke turbine disk in an aero-engine, and can also simulate the cooling process of cold air supplying air to the blades through the double-spoke turbine disk, thereby realizing integrated blade and disk testing experiments, and conducting experimental research on turbine blades and double-spoke turbine disks. Attached Figure Description

[0027] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0028] Figure 1 This is a schematic diagram of the overall structure of the present invention;

[0029] Figure 2 This is a schematic diagram of a turbine blade structure;

[0030] Figure 3 This is a schematic diagram of the cover plate structure;

[0031] Figure 4 This is a schematic diagram of a double-spoke turbine disk structure.

[0032] Figure 5 This is a schematic diagram of the turbine guide vane assembly structure;

[0033] Figure 6 This is a schematic diagram of the rotating shaft structure;

[0034] Among them, 1. Rotor structure; 101. Turbine blade; 102. Double-spoke turbine disk; 103. Cover plate disk; 104. Rotating shaft; 105. Front bearing; 106. Rear bearing; 107. Rear retaining ring of disk; 108. Protrusion; 109. Through hole; 110. Turbine disk rear baffle; 111. Baffle; 112. Second receiving hole; 113. Front air vent; 114. Rear air vent; 115. First air chamber; 116. Second air chamber; 117. First receiving hole; 2. Turbine guide vane assembly; 201. Turbine guide vane; 202. First pre-swirl nozzle; 203. Second pre-swirl nozzle; 204. Third pre-swirl nozzle; 205. Top casing; 3. Front load-bearing casing; 301. Air vent; 4. Rear load-bearing casing. Detailed Implementation

[0035] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0036] Those skilled in the art will understand that the term "comprising" as used in this application means the presence of the stated features, integers, steps, operations, elements, and / or components, but does not exclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof. It should be understood that when we say an element is "connected" or "coupled" to another element, it can be directly connected or coupled to the other element, or there may be intermediate elements present. Furthermore, "connected" or "coupled" as used herein can include wireless connections or wireless coupling. The term "and / or" as used herein includes all or any unit and all combinations of one or more associated listed items.

[0037] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0038] like Figures 1 to 6 As shown, an embodiment of the present invention provides an integrated test structure for an aero-engine bladed disk, comprising:

[0039] A double-spoke turbine disk 102 is mounted on a rotating shaft 104. The double-spoke turbine disk 102 has a disk center and a hollow disk cavity, and the disk center and the hollow disk cavity are connected.

[0040] The turbine blade 101 has a front air intake hole 113 corresponding to the leading edge and the chordal region of the blade, and a rear air intake hole 114 corresponding to the trailing edge of the blade at its bottom. The double-spoke turbine disk 102 has an air chamber at the bottom of the turbine blade 101 and is divided into a first air chamber 115 and a second air chamber 116 by a baffle 111. The bottom of the turbine blade 101 is inserted into the air chamber. The first air chamber 115 corresponds to the front air intake hole 113, and the second air chamber 116 corresponds to the rear air intake hole 114. The first air chamber 115 is connected to the cold air, and the second air chamber 116 is connected to the hollow disk cavity through a through hole 109. The disk cavity is connected to the cold air.

[0041] In this embodiment, the bottom of the turbine blade 101 is installed into the air cavity through a tenon and mortise structure. The partition can be provided with an installation groove at the position corresponding to the bottom of the turbine blade 101, or multiple partitions can be provided to separate the air cavity without affecting the installation of the turbine blade 101.

[0042] In this embodiment, it also includes:

[0043] The cover plate 103 is installed at the front end of the double-spoke turbine disk 102 by threaded screws. The cover plate 103 has a first receiving hole 117. One end of the first receiving hole 117 is connected to the cold air, and the other end is indirectly connected to the first air chamber 115. Multiple first receiving holes 117 can be arranged along the circumference of the cover plate 103. It should be noted that after the turbine blade 101 is installed on the double-spoke turbine disk 102, there is a gap between the bottom of the turbine blade 101 and the tenon groove. After the cold air enters the first receiving hole 117, it can flow into the first air chamber 115 through the aforementioned gap.

[0044] In this embodiment, the double-spoke turbine disk 102 extends axially along the shaft 104 to form a flange edge, which is connected to the disk center. The flange edge is provided with a second receiving hole 112 that communicates with the cold air. Multiple second receiving holes 112 can also be arranged circumferentially along the flange edge.

[0045] In this embodiment, a turbine guide vane assembly 2 is also included, which includes:

[0046] Turbine guide vanes 201 are mounted on the body; a first pre-swirl nozzle 202 is provided on the body corresponding to the edge of the double-spoke turbine disk 102, through which cold air is blown to the edge of the double-spoke turbine disk 102 and seals the front of the double-spoke turbine disk 102; a second pre-swirl nozzle 203 is provided on the body corresponding to the first receiving hole 117, through which cold air is sent into the first receiving hole 117; a third pre-swirl nozzle 204 is provided on the body corresponding to the second receiving hole 112, through which cold air is sent into the second receiving hole 112.

[0047] The casing has an air vent 301 that communicates with the cold air. The casing has an inlet and outlet chamber, which are connected to the air vent 301, the first pre-rotation nozzle 202, the second pre-rotation nozzle 203 and the third pre-rotation nozzle 204 respectively.

[0048] In this embodiment, the direction of the cold air flow is as follows:

[0049] Cold air enters the inlet chamber through the air intake 301 and flows to the first pre-swirl nozzle 202, the second pre-swirl nozzle 203, and the third pre-swirl nozzle 204 respectively. The cold air flowing to the first pre-swirl nozzle 202 is blown to the edge of the double-spoke turbine disk 102 and seals the front of the double-spoke turbine disk 102. The cold air flowing to the second pre-swirl nozzle 203 enters the first receiving hole 117 and then flows to the first air chamber 115 to cool the leading edge and chord region of the blade. The cold air flowing to the third pre-swirl nozzle 204 enters the second receiving hole 112 and then flows to the disk center, enters the hollow disk cavity from the disk center, and then enters the second air chamber 116 through the channel to cool the trailing edge of the blade.

[0050] The three pre-swirl nozzles can cool the double-spoke turbine disk 102 and the turbine blades 101. At the same time, the flow area of ​​the three pre-swirl nozzles can be adjusted to achieve dynamic control of the flow rate of various cooling paths and meet the experimental requirements of different working conditions.

[0051] In this embodiment, the casing includes:

[0052] The front bearing housing 3 is used to support the front bearing 105. The inlet chamber is located in the front bearing housing 3, and the air vent 301 is opened on the front bearing housing 3.

[0053] The top casing 205 is located on the outside of the turbine guide vane 201 and the turbine blade 101, and the top casing 205 is connected to the front casing.

[0054] The rear load-bearing housing 4 is used to support the rear bearing 106 and is connected to the top housing 205; the rotating shaft 104 is rotatably mounted in the front bearing 105 and the rear bearing 106 and is located between the front bearing 105 and the rear bearing 106.

[0055] In this embodiment, the front bearing 105 and the rear bearing 106 are positioned on the rotating shaft 104 by bearing retaining rings.

[0056] In this embodiment, a turbine disk rear baffle 110 is provided at the rear end of the double-spoke turbine disk 102, and the turbine blade 101 is axially fixed to the double-spoke turbine disk 102 by the cover plate 103 and the turbine disk rear baffle 110. It should be noted that the bottom of the turbine blade 101 extends outward by a certain length, and the cover plate 103 and the turbine disk rear baffle 110 can axially position the bottom of the turbine blade 101, thereby achieving axial positioning of the entire turbine blade 101. The specific fixing structure can adopt existing technology, which will not be described in detail here.

[0057] In this embodiment, a thermocouple and a Pitot tube are installed in the inlet chamber; multiple thermocouples are arranged on the front surface and the center surface of the double-spoke turbine disk 102; multiple thermocouples are arranged in the middle of the turbine blade 101; and a flow meter is installed in the bleed port 301. Specifically, the thermocouple located in the middle of the turbine blade 101 is used to measure the temperature of the turbine blade 101 surface; the thermocouples located on the front surface and the center of the double-spoke turbine disk 102 are used to measure the temperature of the front surface and the center of the double-spoke turbine disk 102; the thermocouples and Pitot tube located in the inlet chamber are used to measure temperature and pressure; and the flow meter in the bleed port 301 is used to measure the total flow rate of the cold air.

[0058] In this embodiment, a protrusion 108 and a disc back retaining ring 107 are provided on the rotating shaft 104. The protrusion 108 is located at the front end of the double-spoke turbine disk 102, and the disc back retaining ring 107 is located at the rear end of the double-spoke turbine disk 102. The double-spoke turbine disk 102 is positioned on the rotating shaft 104 by the protrusion 108 and the disc back retaining ring 107.

[0059] In this embodiment, the protrusion 108 has a toothed structure, and cold air can flow from the gap between adjacent protrusions 108 to the center of the disk.

[0060] In this embodiment, the turbine blade 101, the double-spoke turbine disk 102, the cover plate disk 103, the shaft 104, the front bearing 105 and the rear bearing 106 constitute the rotor structure 1, and the turbine guide vane assembly 2 and the casing constitute the stator structure.

[0061] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," "counterclockwise," "axial," "radial," and "circumferential" indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0062] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0063] In this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components, unless otherwise explicitly limited. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0064] The embodiments described above are merely preferred embodiments of the present invention and are not intended to limit the scope of the present invention. Various modifications and improvements made by those skilled in the art to the technical solutions of the present invention without departing from the spirit of the present invention should fall within the protection scope defined by the claims of the present invention.

Claims

1. An integrated test structure for aero-engine bladed disk, characterized in that, include: A double-spoke turbine disk (102) is mounted on a rotating shaft (104). The double-spoke turbine disk (102) has a disk center and a hollow disk cavity, and the disk center and the hollow disk cavity are connected. The turbine blade (101) has a front air intake hole (113) corresponding to the leading edge and the chordal region of the blade, and a rear air intake hole (114) corresponding to the trailing edge of the blade at its bottom. The double-spoke turbine disk (102) is provided with an air chamber at the bottom of the turbine blade (101) and is divided into a first air chamber (115) and a second air chamber (116) by a baffle (111). The bottom of the turbine blade (101) is inserted into the air chamber. The first air chamber (115) corresponds to the front air intake hole (113), and the second air chamber (116) corresponds to the rear air intake hole (114). The first air chamber (115) is connected to the cold air, and the second air chamber (116) is connected to the hollow disk cavity through a through hole (109). The disk cavity is connected to the cold air.

2. The integrated test structure for aero-engine bladed disk according to claim 1, characterized in that, Also includes: The cover plate (103) is installed at the front end of the double-spoke turbine disk (102). The cover plate (103) has a first receiving hole (117). One end of the first receiving hole (117) is connected to the cold air, and the other end is connected to the first air chamber (115).

3. The integrated test structure for aero-engine bladed disk according to claim 2, characterized in that, The double-spoke turbine disk (102) extends axially along the shaft (104) to form a flange edge, which is connected to the disk center. The flange edge is provided with a second receiving hole (112) that is connected to the cold air.

4. The integrated test structure for aero-engine bladed disk according to claim 3, characterized in that, It also includes a turbine guide vane assembly (2), which comprises: Turbine guide vanes (201) are mounted on the body; the body is provided with a first pre-swirl nozzle (202) at the edge of the double-spoke turbine disk (102), through which cold air is blown to the edge of the double-spoke turbine disk (102) and seals the front of the double-spoke turbine disk (102); the body is provided with a second pre-swirl nozzle (203) at the first receiving hole (117), through which cold air is fed into the first receiving hole (117); the body is provided with a third pre-swirl nozzle (204) at the second receiving hole (112), through which cold air is fed into the second receiving hole (112). The casing has an air vent (301) that communicates with cold air. The casing has an inlet and outlet chamber, which are respectively connected to the air vent (301), the first pre-rotation nozzle (202), the second pre-rotation nozzle (203) and the third pre-rotation nozzle (204).

5. The integrated test structure for aero-engine bladed disk according to claim 4, characterized in that, The casing includes: The front bearing housing (3) is used to support the front bearing (105). The inlet chamber is located in the front bearing housing (3). The air vent (301) is opened on the front bearing housing (3). A top casing (205) is disposed on the outside of the turbine guide vanes (201) and turbine blades (101), and the top casing (205) is connected to the front casing; The rear bearing housing (4) is used to support the rear bearing (106) and is connected to the top housing (205); the shaft (104) is installed between the front bearing (105) and the rear bearing (106).

6. The integrated test structure for aero-engine bladed disk according to claim 5, characterized in that, The front bearing (105) and the rear bearing (106) are positioned on the rotating shaft (104) by bearing retaining rings.

7. The integrated test structure for aero-engine bladed disk according to claim 5, characterized in that, The rear end of the double-spoke turbine disk (102) is provided with a turbine disk rear baffle (110), and the turbine blades (101) are axially fixed on the double-spoke turbine disk (102) by the cover plate disk (103) and the turbine disk rear baffle (110).

8. The integrated test structure for aero-engine bladed disk according to claim 5, characterized in that, The inlet chamber is equipped with thermocouples and pitot tubes; the front surface and the center surface of the double-spoke turbine disk (102) are equipped with multiple thermocouples; the middle part of the turbine blade (101) is equipped with multiple thermocouples; and the air vent (301) is equipped with a flow meter.

9. The integrated test structure for aero-engine bladed disk according to claim 1, characterized in that, The rotating shaft (104) is provided with a protrusion (108) and a disc back retaining ring (107). The protrusion (108) is located at the front end of the double-spoke turbine disk (102), and the disc back retaining ring (107) is located at the rear end of the double-spoke turbine disk (102). The double-spoke turbine disk (102) is positioned on the rotating shaft (104) by the protrusion (108) and the disc back retaining ring (107).

10. The integrated test structure for aero-engine bladed disk according to claim 9, characterized in that, The protrusion (108) has a toothed structure, and cold air can flow from the gap between adjacent protrusions (108) to the center of the disk.

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