A disc-edge damping ring brake structure for preventing turbine overspeed in aircraft engines

Abstract

This invention provides a disc-edge damping ring braking structure for preventing turbine overspeed in aero-engines, comprising a disc-edge baffle, a wedge-shaped conical shell, and a damping ring. The disc-edge baffle is installed on the rear side of the turbine disc, the wedge-shaped conical shell is installed on the turbine rear frame, and the damping ring is welded to the wedge-shaped conical shell. In the installed state and under normal engine operation, the wedge-shaped ring surface of the disc-edge baffle shoulder and the wedge-shaped inner ring surface of the damping ring face each other with a certain gap between them. In the event of a shaft breakage failure, the turbine rotor moves backward, and the wedge-shaped ring surface of the disc-edge baffle contacts and rubs against the wedge-shaped ring surface of the damping ring, consuming the kinetic energy of the turbine rotor, limiting the turbine rotor speed, achieving rapid braking, and ensuring the integrity of the turbine rotor and engine safety.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine design, specifically relating to a disc-edge damping ring brake structure for preventing turbine overrun in aero-engines. Background Technology

[0002] During operation, aero-engines may experience shaft breakage due to extreme loads, high-cycle fatigue, corrosion, material defects, manufacturing and assembly errors, or other indirect events. After shaft breakage, the turbine rotor will accelerate under the drive of high-energy combustion gases. If this acceleration reaches a certain level, the turbine disk may rupture due to excessive centrifugal stress. The resulting high-energy fragments could potentially penetrate the engine casing or even the aircraft fuselage, causing catastrophic consequences. Therefore, designing turbine rotor over-spinning protection structures to prevent turbine over-spinning after shaft breakage is an indispensable part of aero-engine design and development.

[0003] Aero engines typically monitor rotor speed using sensors. Upon detecting shaft breakage or turbine disk overspinning, the fuel supply to the combustion chamber is immediately cut off, causing the turbine rotor to decelerate due to the lack of high-energy combustion gas. However, the engine control system still requires time to detect and cut off the fuel supply after a shaft breakage. During this process, the turbine rotor speed rapidly increases. Relying solely on the control system to limit turbine speed after shaft failure remains a very demanding requirement. Therefore, it is necessary to design and install a turbine rotor overspinning protection structure to prevent turbine overspinning after a shaft breakage. Summary of the Invention

[0004] To address the aforementioned technical problems, this invention aims to provide a disc-edge damping ring brake structure for preventing turbine overspeed in aero-engines. This structure can limit turbine speed after a shaft breakage failure, thus preventing catastrophic turbine disc rupture. Furthermore, this damping ring brake structure has the advantages of simple configuration and ease of assembly and disassembly.

[0005] To achieve the above objectives, the present invention adopts the following technical solution:

[0006] A disc-edge damping ring brake structure for preventing over-rotation of an aero-engine turbine includes a disc-edge baffle, a wedge-shaped conical shell, and a damping ring. The disc-edge damping ring brake structure is installed at the location of the low-pressure turbine component and limits the rotational speed of the low-pressure turbine rotor component after the low-pressure shaft breaks, thereby preventing the low-pressure turbine rotor component from over-rotating and the low-pressure turbine disc from breaking.

[0007] The disc rim baffle has a bottom mounting edge, a head baffle, and a shoulder wedge-shaped annular surface; the wedge-shaped conical shell has a bottom mounting edge, a head overlapping edge, and an extended shell; the damping ring has a wedge-shaped outer annular surface and a wedge-shaped inner annular surface;

[0008] The disk edge baffle is connected to the turbine disk mounting edge at the bottom mounting edge by a second bolt to achieve its fixed installation; the disk edge baffle is axially pressed against the rear end face of the tenon structure at the head baffle edge to restrict the axial displacement of the low-pressure turbine blades.

[0009] The wedge-shaped conical shell is fixedly installed by connecting its bottom mounting edge to the head mounting edge of the bearing seat via a third bolt; the overlapping edge of the head of the wedge-shaped conical shell also overlaps with the front end of the flow guide, enhancing the deformation resistance of the wedge-shaped conical shell.

[0010] The damping ring is welded to the outer shell of the wedge-shaped conical shell via a wedge-shaped outer ring surface;

[0011] The shoulder wedge-shaped annular surface of the disc edge baffle and the wedge-shaped inner annular surface of the damping ring are opposite each other in the installed state, and there is a certain gap between them. The distance between the two wedge-shaped annular surfaces can be adjusted by modifying the geometric dimensions of the wedge-shaped conical shell and the thickness of the damping ring.

[0012] Furthermore, during normal operation of the aero-engine, a certain gap is maintained between the shoulder wedge-shaped annular surface of the disc edge baffle and the wedge-shaped inner annular surface of the damping ring, so that no contact friction occurs;

[0013] When the low-pressure shaft breaks, the low-pressure turbine rotor component moves backward, and the disk edge baffle moves backward with the low-pressure turbine disk. The shoulder wedge-shaped ring surface will press against the wedge-shaped inner ring surface, and the two will come into contact and rub against each other, consuming the kinetic energy of the low-pressure turbine rotor component, limiting its speed increase, achieving rapid braking, and thus ensuring the integrity of the low-pressure turbine component.

[0014] Furthermore, it is applied to aero engines with a load-bearing frame behind the last-stage turbine rotor.

[0015] Furthermore, the damping ring uses a metal rubber material, which is a homogeneous elastic porous material made of metal wire.

[0016] The advantages of this invention compared to the prior art are as follows:

[0017] The disc edge damping ring braking structure of the present invention for preventing turbine overrun in aero-engines can quickly limit the speed of the turbine and brake it after a shaft failure in an aero-engine, thereby preventing further catastrophic accidents such as turbine overrun and disc rupture, and thus ensuring the integrity of the turbine.

[0018] The aforementioned disc-edge damping ring brake structure for preventing turbine overrun in aero-engines has the advantages of simple structural configuration, minimal impact on existing engine structures, ease of processing and manufacturing, and low cost.

[0019] The damping ring uses a metal-rubber material, which has a simple structure and good damping effect. Metal-rubber is a homogeneous elastic porous material made of metal wires. It has both high strength characteristics similar to the selected metal material and can generate a damping effect through the internal friction of the metal wires, thus having the function of damping energy dissipation and vibration reduction.

[0020] The aforementioned disc-edge damping ring brake structure for preventing turbine overrun in aero engines can be used not only for low-pressure turbines but also for high-pressure turbines. It can be applied not only to dual-rotor engines but also to aero engines with other structural layouts such as single-rotor or triple-rotor engines. Attached Figure Description

[0021] The above-described and other features, properties, and advantages of the present invention will be more vividly described below in conjunction with the accompanying drawings and examples, wherein:

[0022] Figure 1 This is a schematic diagram of a typical twin-rotor aero-engine.

[0023] Figure 2 This is a partial schematic diagram of a low-pressure turbine component using an example of the present invention.

[0024] Figure 3 This is a partial schematic diagram of a disc rim damping ring brake structure using an example of the present invention. Detailed Implementation

[0025] The present invention will now be described in detail with reference to the accompanying drawings and examples.

[0026] Figure 1 A schematic diagram of a typical twin-rotor aircraft engine 100 is shown. The twin-rotor aircraft engine 100 includes a low-pressure compressor component 11, a high-pressure compressor component 12, a combustion chamber 13, a high-pressure turbine component 14, and a low-pressure turbine component 15, etc.

[0027] The low-pressure compressor component 11 includes a low-pressure compressor rotor component 11A and a low-pressure compressor stator component 11B; the high-pressure compressor component 12 includes a high-pressure compressor rotor component 12A and a high-pressure compressor stator component 12B; the high-pressure turbine component 14 includes a high-pressure turbine rotor component 14A and a high-pressure turbine stator component 14B; and the low-pressure turbine component includes a low-pressure turbine rotor component 15A and a low-pressure turbine stator component 15B.

[0028] The high-pressure compressor rotor assembly 12A is driven by the high-pressure turbine rotor assembly 14A, and the two are connected by a high-pressure shaft 16; the low-pressure compressor rotor assembly 11A is driven by the low-pressure turbine rotor assembly 15A, and the two are connected by a low-pressure shaft 17.

[0029] The low-pressure compressor rotor assembly 11A is supported by the first roller bearing 1 and the first ball bearing 2, and the high-pressure compressor rotor assembly 12A is supported by the second ball bearing 3. The axial and radial forces on these bearings are transmitted outward through the front load-bearing frame 18. The high-pressure turbine rotor assembly 14A is supported by the second roller bearing 4, and the radial force on this bearing is transmitted outward through the high-pressure turbine rear frame 19. The low-pressure turbine rotor assembly 15A is supported by the third roller bearing 5, and the radial force on this bearing is transmitted outward through the low-pressure turbine rear frame 20.

[0030] During the operation of the dual-rotor aero-engine 100, high-temperature, high-energy gas is discharged from the combustion chamber 13 and impacts the high-pressure turbine component 14 and the low-pressure turbine component 15 backward (in the direction of arrow A), driving the high-pressure turbine rotor component 14A and the low-pressure turbine rotor component 15A to rotate. Furthermore, the high-pressure turbine rotor component 14A drives the high-pressure compressor rotor component 12A, and the low-pressure turbine rotor component 15A drives the low-pressure compressor rotor component 11A to rotate.

[0031] When the low-pressure shaft 17 fails, for example, if it breaks at the fracture location 171, on the one hand, the low-pressure turbine rotor component 15A suddenly sheds the load of the low-pressure compressor rotor component 11A, and under the continuous drive of the high-energy gas in the flow channel, the rotational speed increases rapidly; on the other hand, since the fracture location 171 is located behind the first ball bearing 2, the low-pressure shaft 17 loses its axial constraint, and the low-pressure shaft 17 and the low-pressure turbine rotor component 15A will move axially backward (in the direction of arrow A) under the drive of the high-energy gas.

[0032] Figure 2 The disc rim damping ring brake structure 102 of the present invention for preventing turbine overrun is used in the low-pressure turbine component 15 of the aero-engine 101. The aero-engine 101 is structurally similar to that of the dual-rotor aero-engine 100, and the component designations of the dual-rotor aero-engine 100 are used here.

[0033] The low-pressure turbine component 15 includes a low-pressure turbine rotor component 15A and a low-pressure turbine stator component 15B.

[0034] The low-pressure turbine rotor assembly 15A mainly includes a low-pressure turbine disk 21 and low-pressure turbine blades 22. The low-pressure turbine blades 22 are mounted on the low-pressure turbine disk 21 via a tenon joint structure 23. A disk rim baffle 24 and a disk rim baffle 25 are installed before and after the tenon joint structure 23, respectively. The disk rim baffle 24 is mounted on the low-pressure turbine disk 21 via a first bolt 26, and the disk rim baffle 25 is mounted on the low-pressure turbine disk 21 via a second bolt 27.

[0035] The low-pressure turbine stator component 15B mainly includes the bearing housing 31 of the third roller bearing 5 and the low-pressure turbine rear frame 20, wherein the low-pressure turbine rear frame 20 includes the load-bearing spoke 32, the inner ring 33 of the guide vane and the outer ring 34 of the guide vane.

[0036] Figure 3 The specific structure of an example of the present invention and its installation method in the low-pressure turbine component 15 are shown in the figure.

[0037] One embodiment of the present invention includes a disc rim baffle 25, a wedge-shaped conical shell 35, and a damping ring 36.

[0038] The disc edge baffle 25 has a bottom mounting edge 251, a head baffle 252 and a shoulder wedge-shaped annular surface 253; the wedge-shaped conical shell 35 has a bottom mounting edge 351, a head overlapping edge 352 and an extended shell 353; the damping ring 36 has a wedge-shaped outer annular surface 361 and a wedge-shaped inner annular surface 362.

[0039] The disk edge baffle 25 is connected to the turbine disk mounting edge 211 at the bottom mounting edge 251 by the second bolt 27 to achieve its fixed installation; the disk edge baffle 25 is axially pressed with the rear end face 231 of the tenon structure 23 at the head edge 252 to limit the axial displacement of the low-pressure turbine blade 22.

[0040] The wedge-shaped conical shell 35 is connected to the head mounting edge 311 of the bearing housing 31 by a third bolt 36 at the bottom mounting edge 351 position to achieve its fixed installation; the head overlapping edge 352 of the wedge-shaped conical shell 35 also overlaps with the front end 331 of the inner ring 33 of the guide tube to enhance the deformation resistance of the wedge-shaped conical shell 35.

[0041] The damping ring 36 is welded to the outer shell 354 of the wedge-shaped conical shell 25 via the wedge-shaped outer ring surface 362.

[0042] After installation, the shoulder wedge-shaped annular surface 253 of the disc edge baffle 25 and the wedge-shaped inner annular surface 361 of the damping ring 36 are opposite each other, with a certain gap between them. The distance between the two wedge-shaped annular surfaces can be adjusted by modifying the geometric dimensions of the wedge-shaped conical shell 35 and the thickness of the damping ring 36.

[0043] The working principle of this invention is as follows:

[0044] When the aircraft engine 101 is operating normally, a certain gap is maintained between the shoulder wedge-shaped annular surface 253 of the disc edge baffle 25 and the wedge-shaped inner annular surface 361 of the damping ring 36, and no contact friction will occur.

[0045] After the low-pressure shaft 17 breaks, the low-pressure turbine rotor component 15A moves backward, and the disk edge baffle 25 moves backward with the low-pressure turbine disk 21. The shoulder wedge-shaped annular surface 253 will press against the wedge-shaped inner annular surface 361, and the two will come into contact and rub against each other, consuming the kinetic energy of the low-pressure turbine rotor component 15A, limiting its speed increase, achieving rapid braking, and thus ensuring the integrity of the low-pressure turbine component 15.

[0046] The aforementioned disc edge damping ring brake structure 102 for preventing turbine overrun in aero engines has the advantages of simple structural configuration, easy processing and manufacturing, low cost, and minimal impact on the existing engine structure. It can be used not only for low-pressure turbines but also for high-pressure turbines, and can be applied not only to dual-rotor engines but also to aero engines with other structural layouts such as single-rotor or triple-rotor engines.

[0047] The damping ring 36 is made of metal-rubber material, which has a simple structure and good damping effect. Metal-rubber is a homogeneous elastic porous material made of metal wires. It has high strength characteristics similar to the selected metal material, and can also generate a damping effect through the internal friction of the metal wires, thus having the function of damping energy dissipation and vibration reduction.

[0048] The above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit it. Although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art should understand that modifications can still be made to the specific implementation of the present invention or equivalent substitutions can be made to some technical features without departing from the spirit of the technical solutions of the present invention, and all such modifications and substitutions should be covered within the scope of the technical solutions claimed in the present invention.

Claims

1. A disc-edge damping ring brake structure for preventing turbine overrun in aircraft engines, characterized in that: It includes a disc rim baffle, a wedge-shaped conical shell, and a damping ring; the disc rim damping ring brake structure is installed at the location of the low-pressure turbine component to limit the speed of the low-pressure turbine rotor component after the low-pressure shaft breaks, thereby preventing the low-pressure turbine rotor component from over-rotating and the low-pressure turbine disc from breaking. The disc edge baffle has a bottom mounting edge, a head baffle, and a shoulder wedge-shaped annular surface; the wedge-shaped conical shell has a bottom mounting edge, a head overlapping edge, and an extended shell. The damping ring has a wedge-shaped outer ring surface and a wedge-shaped inner ring surface; The disk edge baffle is connected to the turbine disk mounting edge at the bottom mounting edge by a second bolt to achieve its fixed installation; the disk edge baffle is axially pressed against the rear end face of the tenon structure at the head baffle edge to restrict the axial displacement of the low-pressure turbine blades. The wedge-shaped cone shell is fixedly installed by connecting its bottom mounting edge to the head mounting edge of the bearing seat via a third bolt; the head overlapping edge of the wedge-shaped cone shell also overlaps with the front end of the inner ring of the flow guide, enhancing the deformation resistance of the wedge-shaped cone shell. The damping ring is welded to the outer shell of the wedge-shaped conical shell via a wedge-shaped outer ring surface; The shoulder wedge-shaped annular surface of the disc edge baffle and the wedge-shaped inner annular surface of the damping ring are opposite each other in the installed state, and there is a certain gap between them. The distance between the two wedge-shaped annular surfaces can be adjusted by modifying the geometric dimensions of the wedge-shaped conical shell and the thickness of the damping ring. When the aero engine is operating normally, a certain gap is maintained between the shoulder wedge-shaped annular surface of the disc edge baffle and the wedge-shaped inner annular surface of the damping ring, so that no contact friction occurs. When the low-pressure shaft breaks, the low-pressure turbine rotor component moves backward, and the disk edge baffle moves backward with the low-pressure turbine disk. The shoulder wedge-shaped ring surface will press against the wedge-shaped inner ring surface, and the two will come into contact and rub against each other, consuming the kinetic energy of the low-pressure turbine rotor component, limiting its speed increase, achieving rapid braking, and thus ensuring the integrity of the low-pressure turbine component.

2. The disc rim damping ring brake structure for preventing turbine overrunning in an aero-engine according to claim 1, characterized in that: It is used in aircraft engines with a load-bearing frame after the last stage turbine rotor.

3. The disc rim damping ring brake structure for preventing turbine overrunning in an aircraft engine according to claim 1, characterized in that: The damping ring uses a metal rubber material, which is a homogeneous elastic porous material made of metal wire.

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