CN114964787B - Aeroengine complete machine low vortex rotor blade stress measurement structure - Google Patents
Aeroengine complete machine low vortex rotor blade stress measurement structure Download PDFInfo
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- CN114964787B CN114964787B CN202210518372.8A CN202210518372A CN114964787B CN 114964787 B CN114964787 B CN 114964787B CN 202210518372 A CN202210518372 A CN 202210518372A CN 114964787 B CN114964787 B CN 114964787B
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/02—Details or accessories of testing apparatus
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01L—MEASURING FORCE, STRESS, TORQUE, WORK, MECHANICAL POWER, MECHANICAL EFFICIENCY, OR FLUID PRESSURE
- G01L1/00—Measuring force or stress, in general
- G01L1/08—Measuring force or stress, in general by the use of counterbalancing forces
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- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
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Abstract
The application provides a stress measurement structure of a complete machine low-vortex rotor blade of an aeroengine, which is used for leading out a lead of a lead-out device to realize the acquisition of stress level and distribution measurement signals on a final low-pressure turbine rotor blade, wherein the lead-out device is designed to cover the lead-out device by a heat shield, the lead-out device has a protection function in a high-temperature environment, cooling water can be introduced into an annular cooling channel in the heat shield through a water inlet pipeline, and can flow out through a gap between a water outlet pipeline and an air charging pipeline, so that heat in the high-temperature environment is absorbed and taken away, the heat in the heat shield is kept at a lower temperature, the lead-out device can be effectively protected from high-temperature damage, the air charging pipeline is kept at a low temperature, the lead-out device led out from the air charging pipeline is protected from high-temperature damage, meanwhile, the air can be charged into the heat shield through the air charging pipeline, a certain pressure is kept in the heat shield, and the air in the high-temperature environment is prevented from directly entering the heat shield to impact the lead-out device, and damage to the lead-out device.
Description
Technical Field
The application belongs to the technical field of low-vortex rotor blade stress measurement under the complete machine condition of an aeroengine, and particularly relates to a complete machine low-vortex rotor blade stress measurement structure of the aeroengine, which is used for leading out a lead of a final-stage low-pressure turbine rotor blade stress measurement strain gauge.
Background
The aeroengine low pressure turbine tail structure includes:
a low pressure turbine outer casing 1;
the final low-pressure turbine rotor disk 2 is arranged in the low-pressure turbine outer casing 1 and is positioned in the low-pressure turbine outer casing 1;
the last-stage low-pressure turbine rotor blade 3 is arranged in the low-pressure turbine outer casing 1, and the blade root part is connected to the outer edge of the low-pressure turbine rotor disk 2 along the circumferential direction;
the low-pressure turbine inner-layer casing 4 is arranged in the low-pressure turbine outer-layer casing 1 and positioned at the rear end of the low-pressure turbine outer-layer casing 1, and a flow passage is formed between the low-pressure turbine inner-layer casing 4 and the low-pressure turbine outer-layer casing 1;
the low-pressure turbine hollow rotating shaft 5 penetrates through the disc center of the final-stage low-pressure turbine rotor disc 2, and the outer wall of the rear end is connected with the side wall of the final-stage low-pressure turbine rotor disc 2 facing the rear end of the low-pressure turbine outer casing 1 through an annular connecting edge;
the rear end support bearing 6 is sleeved at the rear end of the hollow rotating shaft 5 of the low-pressure turbine;
the bearing mounting seat 7 is connected in the low-pressure turbine inner-layer casing 4 and is provided with a bearing mounting hole; the rear end support bearing 6 is arranged in the bearing mounting hole;
the axle center oil collecting ring 8 is connected to the rear end of the low-pressure turbine hollow rotating shaft 5 and is in interference fit with the low-pressure turbine hollow rotating shaft 5;
the stress application diffuser outer casing 9 is in butt joint with the rear end of the low-pressure turbine outer casing 1;
the inner cone 10 is positioned on the inner side of the outer casing 9 of the stress diffuser and is butted with the rear end of the inner casing 4 of the low-pressure turbine.
The final low-pressure turbine rotor blade 3 at the tail of the low-pressure turbine of the aero-engine bears larger complex stress in the working process of the aero-engine, and the stress level and the distribution thereof are accurately measured, so that the method has great significance for the design and improvement of the aero-engine.
Currently, the stress level and the distribution of the stress level on the final low-pressure turbine rotor blade 3 are measured by a strain gauge 11 arranged on the final low-pressure turbine rotor blade 3, and the lead wires of the strain gauge 11 are led out from the inside of the low-pressure turbine hollow rotating shaft 5 through an axle center oil collecting ring 8 along the outer wall of the final low-pressure turbine rotor blade 3, the outer wall of the final low-pressure turbine rotor disk 2 and the lead wire hole penetrating through the outer wall of the rear end of the low-pressure turbine hollow rotating shaft 5, and are connected to a lead wire 12 arranged in an inner cone 10, and the lead wire of the lead wire 12 is led out through the lead wire holes arranged on the inner cone 10 and a stress diffuser outer casing 9, so that the stress level and the distribution measurement signals of the stress level on the final low-pressure turbine rotor blade 3 are collected, and the technical scheme has the following defects:
1) The tail structure of the low-pressure turbine of the aero-engine is damaged, the possibility of pressure leakage is generated, the overall performance of the aero-engine is affected, and accurate measurement results of the stress level and the distribution of the stress level on the final low-pressure turbine rotor blade 3 are difficult to obtain;
2) The electrical conductor 12 is exposed to high temperature conditions, and is subject to performance damage.
The present application has been made in view of the above-described technical drawbacks.
It should be noted that the above disclosure of the background art is only for aiding in understanding the inventive concept and technical solution of the present application, which is not necessarily prior art to the present patent application, and should not be used for evaluating the novelty and creativity of the present application in the case where no clear evidence indicates that the above content has been disclosed at the filing date of the present application.
Disclosure of Invention
The application aims to provide a stress measuring structure of a complete machine low-vortex rotor blade of an aeroengine, which is used for leading out a lead of a strain gauge for measuring the stress of a final stage low-pressure turbine rotor blade so as to overcome or alleviate at least one technical defect of the prior art.
The technical scheme of the application is as follows:
an aeroengine complete machine low vortex rotor blade stress measurement structure, comprising:
the support plate is arranged in the inner cone and connected to the bearing mounting seat, and is provided with a perforation;
the electric lead is arranged in the inner cone, the stator part of the electric lead is connected to the supporting plate, the rotor part of the electric lead passes through the perforation to be connected with the axle center oil collecting ring, and the electric lead is connected to the strain gauge;
the heat shield is arranged in the inner cone and connected to the supporting plate to shield the electric guide device, an annular cooling channel is arranged in the heat shield, a heat shield water inlet hole and a heat shield water outlet hole which are communicated with the annular cooling channel are formed in the outer wall of the heat shield, and a heat shield lead hole which is communicated with the annular cooling channel is formed in the inner wall of the heat shield;
the outlet end of the water inlet pipeline passes through the water inlet holes on the outer casing and the inner cone of the stress application diffuser and is communicated with the water inlet holes of the heat shield;
the inlet end of the water outlet pipeline passes through the water outlet holes on the outer casing and the inner cone of the diffuser and is communicated with the water outlet holes of the heat shield;
the outlet end of the air charging pipeline penetrates through the water outlet pipeline, the water outlet hole of the heat shield and the annular cooling channel, and is communicated with the lead hole of the heat shield, and leads of the lead device are led out.
According to at least one embodiment of the present application, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the structure further includes:
the water inlet hose is connected between the outlet end of the water inlet pipeline and the water inlet hole of the heat shield;
the water outlet hose is connected between the inlet end of the water outlet pipeline and the water outlet hole of the heat shield.
According to at least one embodiment of the application, in the stress measuring structure of the low-vortex rotor blade of the whole aeroengine, a plurality of water inlet pipes, a plurality of water inlet holes on the outer casing of the stress diffuser and the inner cone, a plurality of water inlet holes on the heat shield and a plurality of water inlet hoses are distributed along the circumference of the heat shield.
According to at least one embodiment of the present application, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the structure further includes:
the heat insulation pipeline is arranged in the water outlet pipeline, sleeved on the periphery of the air inflation pipeline and filled with heat insulation materials between the air inflation pipeline and the heat insulation pipeline.
According to at least one embodiment of the application, in the stress measuring structure of the low-vortex rotor blade of the whole aeroengine, a bearing mounting seat is provided with a cooling air inlet, and the cooling air inlet is communicated with the space among the supporting plate, the heat shield and the inner cone;
the inner cone is provided with a cooling air outlet hole which is communicated with the space among the supporting plate, the heat shield and the inner cone.
According to at least one embodiment of the present application, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the structure further includes:
the outlet end of the air supply pipeline passes through air supply holes on the low-pressure turbine outer casing and the low-pressure turbine inner casing and is communicated with the space between the bearing mounting seat and the final low-pressure turbine rotor disk;
the annular cover is sleeved on the periphery of the axle center oil collecting ring and is connected with the supporting plate, and an air supply channel is formed between the annular cover and the supporting plate; the air supply channel is communicated with the space between the bearing mounting seat and the final-stage low-pressure turbine rotor disk through an air supply hole on the bearing mounting seat, and is communicated with the axle center oil collecting ring through an air outlet hole on the axle center oil collecting ring.
According to at least one embodiment of the present application, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the structure further includes:
and the sleeve shaft is connected between the rotor part of the lead electric appliance and the axle center oil collecting ring, and the lead of the hollow supply transformer passes through the sleeve shaft.
The application has at least the following beneficial technical effects:
the utility model provides a low vortex rotor blade stress measurement structure of aeroengine complete machine for draw forth the lead wire of electrical apparatus, realize the collection to stress level and distribution measurement signal on final stage low pressure turbine rotor blade, wherein design covers the electrical apparatus with the heat exchanger, have the guard action to the electrical apparatus under high temperature environment, and can pass through the inlet channel and let in the annular cooling channel in the heat exchanger, the cooling water can flow out through outlet channel, the gap between the inflation pipeline, thereby absorb and take away the heat of high temperature environment, make the heat exchanger keep lower temperature in the heat exchanger cover, thereby can effectively protect the electrical apparatus not suffer high temperature damage, and make the inflation pipeline keep low temperature, the electrical apparatus lead wire of leading forth in the protection is not suffered high temperature damage through the inflation pipeline, simultaneously, can fill into low temperature high pressure gas in the heat exchanger through the inflation pipeline, make the heat exchanger keep certain pressure, avoid the interior gas of high temperature environment directly to get into the heat exchanger in the impact electrical apparatus, cause the damage to the electrical apparatus.
Drawings
FIG. 1 is a schematic illustration of a last stage low pressure turbine rotor blade strain gauge lead out structure provided by an embodiment of the present application;
FIG. 2 is a partial cross-sectional view of a last stage low pressure turbine rotor blade strain gauge lead out structure provided by an embodiment of the present application;
wherein:
1-a low-pressure turbine outer casing; 2-final stage low pressure turbine rotor disk; 3-last stage low pressure turbine rotor blades; 4-a low-pressure turbine inner casing; 5-a hollow shaft of a low-pressure turbine; 6-a rear end support bearing; 7-a bearing mounting seat; 8-an axle center oil collecting ring; 9-stressing the diffuser outer casing; 10-an inner cone; 11-strain gauge; 12-an electric primer; 13-a support plate; 14-a heat shield; 15-a water inlet pipeline; 16-a water outlet pipeline; 17-an inflation pipeline; 18-a water inlet hose; 19-a water outlet hose; 20-insulating tubing; 21-an air supply pipeline; 22-an annular cover; 23-quill.
For the purpose of better illustrating the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced and do not represent the actual product dimensions, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present patent.
Detailed Description
In order to make the technical solution of the present application and its advantages more clear, the technical solution of the present application will be further and completely described in detail with reference to the accompanying drawings, it being understood that the specific embodiments described herein are only some of the embodiments of the present application, which are for explanation of the present application and not for limitation of the present application. It should be noted that, for convenience of description, only the part related to the present application is shown in the drawings, and other related parts may refer to the general design, and the embodiments of the present application and the technical features of the embodiments may be combined with each other to obtain new embodiments without conflict.
Furthermore, unless defined otherwise, technical or scientific terms used in the description of the application should be given the ordinary meaning as understood by one of ordinary skill in the art to which the application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," "outer," and the like as used in the description of the present application are merely used for indicating relative directions or positional relationships, and do not imply that the devices or elements must have a specific orientation, be constructed and operated in a specific orientation, and that the relative positional relationships may be changed when the absolute position of the object to be described is changed, thus not being construed as limiting the application. The terms "first," "second," "third," and the like, as used in the description of the present application, are used for descriptive purposes only and are not to be construed as indicating or implying any particular importance to the various components. The use of the terms "a," "an," or "the" and similar referents in the description of the application are not to be construed as limiting the amount absolutely, but rather as existence of at least one. As used in this description of the application, the terms "comprises," "comprising," or the like are intended to cover an element or article that appears before the term as such, but does not exclude other elements or articles from the list of elements or articles that appear after the term.
Furthermore, unless specifically stated and limited otherwise, the terms "mounted," "connected," and the like in the description of the present application are used in a broad sense, and for example, the connection may be a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can also be communicated with the inside of two elements, and the specific meaning of the two elements can be understood by a person skilled in the art according to specific situations.
The application is described in further detail below with reference to fig. 1 to 2.
An aeroengine complete machine low vortex rotor blade stress measurement structure, comprising:
a support plate 13 provided in the inner cone 10 and connected to the bearing mount 7, and having a through hole formed thereon;
a lead 12 provided in the inner cone 10, a stator part of which is connected to the support plate 13, a rotor part of which is connected to the axial oil collecting ring 8 through a through hole, and a lead wire connected to the strain gauge 11;
a heat shield 14 provided in the inner cone 10 and connected to the support plate 13 to shield the current collector 12, wherein the inner wall of the heat shield is provided with an annular cooling channel, the outer wall of the heat shield is provided with a heat shield water inlet hole and a heat shield water outlet hole communicated with the annular cooling channel, and the inner wall of the heat shield is provided with a heat shield lead hole communicated with the annular cooling channel;
the outlet end of the water inlet pipeline 15 passes through water inlet holes on the outer casing 9 and the inner cone 10 of the stress application diffuser and is communicated with the water inlet holes of the heat shield;
a water outlet pipeline 16, the inlet end of which passes through water outlet holes on the outer casing 9 and the inner cone 10 of the diffuser and is communicated with the water outlet holes of the heat shield;
the outlet end of the air charging pipeline 17 passes through the water outlet pipeline 16, the water outlet hole of the heat shield and the annular cooling channel, and is communicated with the lead hole of the heat shield, and leads of the lead electric device 12 are led out.
When the aeroengine works, the stress measuring structure of the whole low-vortex rotor blade of the aeroengine disclosed by the embodiment can lead out the lead wire of the electric guide 12 to realize the acquisition of the stress level and the distribution measuring signal of the stress level on the final low-pressure turbine rotor blade 3, wherein the electric guide 12 is covered by the heat insulation cover 14, the electric guide 12 is protected in a high-temperature environment, cooling water can be introduced into the annular cooling channel in the heat insulation cover 14 through the water inlet pipeline 15, and can flow out through the gap between the water outlet pipeline 16 and the air charging pipeline 17, so that the heat of the high-temperature environment is absorbed and taken away, the heat in the heat insulation cover 14 is kept at a lower temperature, the electric guide 12 can be effectively protected from being damaged at a high temperature, the air charging pipeline 17 is kept at a low temperature, the lead wire of the electric guide 12 led out from being damaged at a high temperature, meanwhile, the air charging low-temperature high-pressure gas into the heat insulation cover 14 can be filled into the heat insulation cover 14 through the air charging pipeline 17, a certain pressure is maintained in the heat insulation cover 14, and the air in the heat insulation cover 14 is prevented from being directly introduced into the heat insulation cover 14 to impact the electric guide 12, and damaging the electric guide 12.
The stress measuring structure of the whole low-vortex rotor blade of the aeroengine disclosed by the embodiment can be similarly applied to the extraction of the temperature and stress measuring sensor leads of the final low-pressure turbine rotating part of the aeroengine.
In some optional embodiments, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the method further includes:
the water inlet hose 18 is connected between the outlet end of the water inlet pipeline 15 and the water inlet hole of the heat shield, and is flexibly connected between the outlet end of the water inlet pipeline 15 and the water inlet hole of the heat shield so as to coordinate the deformation among the outer casing 9, the inner cone 10 and the heat shield 14 of the stress diffuser and avoid generating larger local stress;
the water outlet hose 19 is connected between the inlet end of the water outlet pipeline 16 and the water outlet hole of the heat shield, and is flexibly connected between the inlet end of the water outlet pipeline 16 and the water outlet hole of the heat shield so as to coordinate the deformation among the outer casing 9 of the stress diffuser, the inner cone 10 and the heat shield 14 and avoid generating larger local stress.
In some alternative embodiments, in the stress measurement structure of the low vortex rotor blade of the whole aeroengine, the water inlet pipeline 15 and the corresponding stress application diffuser outer casing 9 and the water inlet holes, the heat shield water inlet holes and the water inlet hoses 18 on the inner cone 10 are multiple and distributed along the circumference of the heat shield 14.
In some optional embodiments, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the method further includes:
the heat insulating pipe 20 is provided in the water outlet pipe 16, and is fitted around the outer periphery of the air charging pipe 17, and a heat insulating material is filled between the air charging pipe 17 and the heat insulating pipe.
For the stress measurement structure of the whole low-vortex rotor blade of the aeroengine disclosed in the above embodiment, it can be understood by those skilled in the art that the cooling water introduced into the annular cooling channel in the heat shield 14 can flow out through the gaps between the water outlet pipeline 16 and the heat insulation pipeline 20, and the heat insulation material is filled between the heat insulation pipeline 20 and the air charging pipeline 17, so that heat transfer into the air charging pipeline 17 can be avoided, and the air charging pipeline 17 can keep a lower temperature under the condition of introducing gas into the air charging pipeline 17.
In some alternative embodiments, in the stress measuring structure of the low-vortex rotor blade of the whole aeroengine, the bearing mounting seat 7 is provided with a cooling air inlet, and the cooling air inlet is communicated with the space among the supporting plate 13, the heat shield 14 and the inner cone 10;
the inner cone 10 is provided with a cooling air outlet hole which is communicated with the space among the support plate, the heat shield 14 and the inner cone 10.
For the stress measurement structure of the low-vortex rotor blade of the whole aeroengine disclosed in the above embodiment, it can be understood by those skilled in the art that the cold air in the inner casing 4 of the low-pressure turbine can enter the space between the heat shield 14 and the inner cone 10 through the cooling air inlet of the bearing mounting seat, absorb and take away the heat in the space, further ensure that the lower temperature is maintained in the heat shield 14, and protect the electric guider 12 from high-temperature damage.
In some optional embodiments, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the method further includes:
the outlet end of the air supply pipeline 21 passes through air supply holes on the low-pressure turbine outer casing 1 and the low-pressure turbine inner casing 4 and is communicated with the space between the bearing mounting seat 7 and the final low-pressure turbine rotor disk 2;
the annular cover 22 is sleeved on the periphery of the axle center oil collecting ring 8, is connected with the supporting plate 13 and forms an air supply channel with the supporting plate 13; the air supply channel is communicated with the space between the bearing mounting seat 7 and the final-stage low-pressure turbine rotor disk 2 through an air supply hole on the bearing mounting seat 7, and is communicated with the inside of the axle center oil collecting ring 8 through an air outlet hole on the axle center oil collecting ring 8.
For the stress measuring structure of the whole low-vortex rotor blade of the aeroengine disclosed by the embodiment, as can be understood by those skilled in the art, cooling air can be introduced into the space between the bearing mounting seat 7 and the final low-pressure turbine rotor disk 2 through the air supply pipeline 21, the cooling air can enter the air supply channel formed between the annular cover 22 and the supporting plate 13 through the air supply hole on the bearing mounting seat 7, then flows out of the axle center oil collecting ring 8 and the low-pressure turbine hollow rotating shaft 5 through the air outlet hole on the axle center oil collecting ring 8, absorbs and takes away heat in the air supply channel, cools the rotor component of the electric guide 12, avoids the rotor component of the electric guide 12 from being damaged by high temperature, can maintain a certain pressure in the air supply channel, avoids the air entering in the high-temperature environment, impacts the rotor component of the electric guide 12, damages the rotor component, and simultaneously the flow channel can realize the sealing function of the sliding oil cavity.
In some optional embodiments, in the stress measurement structure of the low-vortex rotor blade of the whole aeroengine, the method further includes:
the sleeve shaft 23 is connected between the rotor component of the electric lead 12 and the axle center oil collecting ring 8, the lead wire of the hollow supply transformer 11 passes through the sleeve shaft, effective torque transmission is carried out between the electric lead 12 and the low-pressure turbine hollow rotating shaft 5, the two ends of the sleeve shaft can be designed to have arc structures, the sleeve shaft can be used for compensating different axiality between the electric lead 12 and the low-pressure turbine hollow rotating shaft 5, the measurement precision is ensured, and in addition, the sleeve shaft 23 can be broken and disconnected in time when in failure, so that the electric lead 12 is protected from mechanical damage.
In the description, each embodiment is described in a progressive manner, and each embodiment is mainly described by the differences from other embodiments, so that the same similar parts among the embodiments are mutually referred.
Having thus described the technical aspects of the present application with reference to the preferred embodiments shown in the drawings, it should be understood by those skilled in the art that the scope of the present application is not limited to the specific embodiments, and those skilled in the art may make equivalent changes or substitutions to the related technical features without departing from the principle of the present application, and those changes or substitutions will fall within the scope of the present application.
Claims (7)
1. An aeroengine complete machine low vortex rotor blade stress measurement structure for draw forth last stage low pressure turbine rotor blade stress measurement strain gauge's lead wire, characterized in that includes:
a support plate (13) arranged in the inner cone (10) and connected to the bearing mounting seat (7) and provided with a perforation;
a lead (12) arranged in the inner cone (10), a stator part of the lead is connected to the supporting plate (13), a rotor part of the lead is connected with the axle center oil collecting ring (8) through the perforation, and a lead connected with the strain gauge (11);
a heat shield (14) arranged in the inner cone (10) and connected to the supporting plate (13) to cover the electric guide (12), wherein an annular cooling channel is arranged in the heat shield, a heat shield water inlet hole and a heat shield water outlet hole which are communicated with the annular cooling channel are arranged on the outer wall of the heat shield, and a heat shield lead hole which is communicated with the annular cooling channel is arranged on the inner wall of the heat shield;
the outlet end of the water inlet pipeline (15) passes through water inlet holes on the outer casing (9) and the inner cone (10) of the stress application diffuser and is communicated with the water inlet holes of the heat shield;
a water outlet pipeline (16), the inlet end of which passes through water outlet holes on the outer casing (9) and the inner cone (10) of the diffuser and is communicated with the water outlet holes of the heat shield;
and the outlet end of the air charging pipeline (17) passes through the water outlet pipeline (16), the heat shield water outlet hole and the annular cooling channel, is communicated with the heat shield lead hole, and leads out the lead of the lead device (12).
2. The aeroengine complete machine low vortex rotor blade stress measurement structure of claim 1, wherein,
further comprises:
a water inlet hose (18) connected between the outlet end of the water inlet pipeline (15) and the water inlet hole of the heat shield;
and the water outlet hose (19) is connected between the inlet end of the water outlet pipeline (16) and the water outlet hole of the heat shield.
3. The aeroengine complete machine low vortex rotor blade stress measurement structure of claim 1, wherein,
the water inlet pipeline (15) and the corresponding water inlet holes, the heat shield water inlet holes and the water inlet hoses (18) on the outer casing (9) and the inner cone (10) of the stress application diffuser are distributed along the circumference of the heat shield (14).
4. The aeroengine complete machine low vortex rotor blade stress measurement structure of claim 1, wherein,
further comprises:
and the heat insulation pipeline (20) is arranged in the water outlet pipeline (16), sleeved on the periphery of the air inflation pipeline (17), and a heat insulation material is filled between the air inflation pipeline (17).
5. The aeroengine complete machine low vortex rotor blade stress measurement structure of claim 1, wherein,
the bearing mounting seat (7) is provided with a cooling air inlet hole, and the cooling air inlet hole is communicated with the space among the supporting plate (13), the heat shield (14) and the inner cone (10);
the inner cone (10) is provided with a cooling air outlet hole, and the cooling air outlet hole is communicated with the space among the supporting plate (13), the heat shield (14) and the inner cone (10).
6. The aeroengine complete machine low vortex rotor blade stress measurement structure of claim 1, wherein,
further comprises:
the outlet end of the air supply pipeline (21) passes through air supply holes on the low-pressure turbine outer casing (1) and the low-pressure turbine inner casing (4) and is communicated with the space between the bearing mounting seat (7) and the final low-pressure turbine rotor disk (2);
the annular cover (22) is sleeved on the periphery of the axle center oil collecting ring (8), is connected with the supporting plate (13) and forms an air supply channel with the supporting plate (13); the air supply channel is communicated with the space between the bearing mounting seat (7) and the final-stage low-pressure turbine rotor disk (2) through an air supply hole on the bearing mounting seat (7), and is communicated with the inside of the axle center oil collecting ring (8) through an air outlet hole on the axle center oil collecting ring (8).
7. The aeroengine complete machine low vortex rotor blade stress measurement structure of claim 1, wherein,
further comprises:
and a sleeve shaft (23) connected between the rotor component of the lead-in device (12) and the shaft center oil collecting ring (8), wherein the lead of the hollow supply transformer (11) passes through.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
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| CN202210518372.8A CN114964787B (en) | 2022-05-12 | 2022-05-12 | Aeroengine complete machine low vortex rotor blade stress measurement structure |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202210518372.8A CN114964787B (en) | 2022-05-12 | 2022-05-12 | Aeroengine complete machine low vortex rotor blade stress measurement structure |
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| CN114964787A CN114964787A (en) | 2022-08-30 |
| CN114964787B true CN114964787B (en) | 2023-09-22 |
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| CN115342774B (en) * | 2022-10-20 | 2023-05-12 | 北京航天动力研究所 | Strain measurement system for high-speed flexible rotor turbine disk of liquid rocket engine |
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