CN114964787A - Stress measurement structure for whole-machine low-vortex rotor blade of aero-engine - Google Patents
Stress measurement structure for whole-machine low-vortex rotor blade of aero-engine Download PDFInfo
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- CN114964787A CN114964787A CN202210518372.8A CN202210518372A CN114964787A CN 114964787 A CN114964787 A CN 114964787A CN 202210518372 A CN202210518372 A CN 202210518372A CN 114964787 A CN114964787 A CN 114964787A
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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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Abstract
The application provides a stress measurement structure of a whole aircraft low-vortex rotor blade of an aeroengine, a lead of an electrical leading device is led out, the acquisition of the stress level and the distribution measurement signal on a final stage low-pressure turbine rotor blade is realized, wherein the electrical leading device is covered by a heat insulation cover, the electrical leading device has a protection effect in a high-temperature environment, cooling water can be introduced into an annular cooling channel in the heat insulation cover through a water inlet pipeline, the cooling water can flow out through a gap between a water outlet pipeline and an inflation pipeline, so that the heat of the high-temperature environment is absorbed and taken away, the temperature in the heat insulation cover is kept at a lower temperature, the electrical leading device can be effectively protected from being damaged by the high temperature, the low temperature in the inflation pipeline is kept, the lead of the electrical leading device led out from the inflation pipeline is protected from being damaged by the high temperature, meanwhile, gas can be filled into the heat insulation cover through the inflation pipeline, so that a certain pressure is kept in the heat insulation cover, the gas in the high-temperature environment is prevented from directly entering the heat shield to impact the electric leading device, and the electric leading device is prevented from being damaged.
Description
Technical Field
The application belongs to the technical field of stress measurement of low-vortex rotor blades under the condition of an aircraft engine complete machine, and particularly relates to a stress measurement structure of a low-vortex rotor blade of an aircraft engine complete machine, which is used for leading out a lead of a stress measurement strain gauge of a last-stage low-pressure turbine rotor blade.
Background
Aircraft engine low pressure turbine afterbody structure includes:
a low-pressure turbine outer casing 1;
the final-stage low-pressure turbine rotor disc 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 disc 2 along the circumferential direction;
the low-pressure turbine inner-layer casing 4 is arranged in the low-pressure turbine outer-layer casing 1, is positioned at the rear end of the low-pressure turbine outer-layer casing 1, and forms a flow channel with the low-pressure turbine outer-layer casing 1;
the low-pressure turbine hollow rotating shaft 5 penetrates through the center of the last-stage low-pressure turbine rotor disc 2, and the outer wall of the rear end of the low-pressure turbine hollow rotating shaft is connected with the side wall of the last-stage low-pressure turbine rotor disc 2, which faces 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 low-pressure turbine hollow rotating shaft 5;
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 axis 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 thrust augmentation diffuser outer casing 9 is butted at the rear end of the low-pressure turbine outer casing 1;
and the inner cone 10 is positioned at the inner side of the thrust augmentation diffuser outer casing 9 and is butted at the rear end of the low-pressure turbine inner-layer casing 4.
The last stage low pressure turbine rotor blade 3 at the tail part of the low pressure turbine of the aeroengine bears larger complex stress in the working process of the aeroengine, the stress level and the distribution thereof are accurately measured, and the method has important significance for the design and the improvement of the aeroengine.
At present, the strain gauge 11 arranged on the last stage low pressure turbine rotor blade 3 is mostly used for measuring the stress level and the distribution on the last stage low pressure turbine rotor blade 3, the lead of the strain gauge 11 is led out from the inside of the low pressure turbine hollow rotating shaft 5 through the axle center oil receiving ring 8 along the outer wall of the last stage low pressure turbine rotor blade 3, the outer wall of the last stage low pressure turbine rotor disk 2 and the lead hole penetrating through the outer wall of the rear end of the low pressure turbine hollow rotating shaft 5, and is connected to the electrical lead 12 arranged in the inner cone 10, and the lead of the electrical lead 12 is led out through the lead holes arranged on the inner cone 10 and the stress application diffuser outer casing 9, so as to realize the acquisition of the stress level and the distribution measurement signal on the last stage low pressure turbine rotor blade 3, and the following defects exist in the technical scheme:
1) the tail structure of the low-pressure turbine of the aircraft engine is damaged, the possibility of pressure leakage is generated, the overall performance of the aircraft engine is influenced, and the accurate measurement result of the stress level and the distribution of the stress level on the last-stage low-pressure turbine rotor blade 3 is difficult to obtain;
2) the electrical starter 12 is in a high temperature environment, and the performance is damaged and is easily damaged.
The present application has been made in view of the above-mentioned technical drawbacks.
It should be noted that the above background disclosure is only for the purpose of assisting understanding of the inventive concept and technical solutions of the present invention, and does not necessarily belong to the prior art of the present patent application, and the above background disclosure should not be used for evaluating the novelty and inventive step of the present application without explicit evidence to suggest that the above content is already disclosed at the filing date of the present application.
Disclosure of Invention
The application aims to provide a stress measurement structure of a whole low-pressure turbine rotor blade of an aircraft engine, which is used for leading out a lead of a stress measurement strain gauge of a last-stage low-pressure turbine rotor blade so as to overcome or alleviate at least one technical defect in the prior art.
The technical scheme of the application is as follows:
a whole-aircraft engine low-turbine rotor blade stress measurement structure comprises:
the supporting plate is arranged in the inner cone, is connected to the bearing mounting seat and is provided with a through hole;
the stator part of the electric initiator is connected to the supporting plate, and the rotor part of the electric initiator passes through the through hole to be connected with the axis oil collecting ring and is connected to a lead of the strain gauge;
the heat shield is arranged in the inner cone, connected to the support plate and used for shielding the electric initiator, 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 penetrates through water inlet holes in the outer casing and the inner cone of the booster diffuser and is communicated with the water inlet hole of the heat insulation cover;
the inlet end of the water outlet pipeline penetrates through water outlet holes in the outer casing and the inner cone of the diffuser and is communicated with the water outlet hole of the heat insulation cover;
and 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, is communicated with the lead hole of the heat shield, and leads of the electric lead are led out from the air charging pipeline.
According to at least one embodiment of the application, the stress measurement structure for the whole low-turbine rotor blade of the aircraft engine further comprises:
the water inlet hose is connected between the outlet end of the water inlet pipeline and the water inlet hole of the heat insulation cover;
and the water outlet hose is connected between the inlet end of the water outlet pipeline and the water outlet hole of the heat insulation cover.
According to at least one embodiment of the application, in the stress measurement structure of the whole low-turbine rotor blade of the aircraft engine, the water inlet pipe and the corresponding booster diffuser outer casing as well as the water inlet holes, the heat shield water inlet holes and the water inlet hoses on the inner cone are multiple and distributed along the circumferential direction of the heat shield.
According to at least one embodiment of the application, the stress measurement structure for the whole low-turbine rotor blade of the aircraft engine further comprises:
and the heat insulation pipeline is arranged in the water outlet pipeline, is sleeved on the periphery of the inflation pipeline, and is filled with heat insulation materials between the inflation pipeline and the heat insulation pipeline.
According to at least one embodiment of the application, in the stress measurement structure of the whole low-turbine rotor blade of the aircraft engine, the bearing mounting seat is provided with a cooling air inlet hole, and the cooling air inlet hole is communicated with spaces among the support plate, the heat shield and the inner cone;
the inner cone is provided with a cooling air outlet which is communicated with the space among the support plate, the heat shield and the inner cone.
According to at least one embodiment of the application, the stress measurement structure for the whole low-turbine rotor blade of the aircraft engine further comprises:
the outlet end of the gas supply pipeline penetrates through gas supply holes in 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-stage low-pressure turbine rotor disc;
the annular cover is sleeved on the periphery of the axis oil collecting ring, is connected with the supporting plate and forms an air supply channel with the supporting plate; the ventilation channel is communicated with the space between the bearing mounting seat and the final-stage low-pressure turbine rotor disc through an air supply hole in the bearing mounting seat and is communicated with the axis oil collecting ring through an air outlet hole in the axis oil collecting ring.
According to at least one embodiment of the application, the stress measurement structure for the whole low-turbine rotor blade of the aircraft engine further comprises:
the sleeve shaft is connected between the rotor component of the electrical lead and the axis oil collecting ring, and a hollow lead of the supply transformer passes through the sleeve shaft.
The application has at least the following beneficial technical effects:
the stress measuring structure of the whole low-vortex rotor blade of the aircraft engine is used for leading out a lead of an electrical lead so as to realize the acquisition of the stress level and the distribution measuring signal of the final-stage low-pressure turbine rotor blade, wherein the electrical lead is covered by a heat shield, the electrical lead is protected in a high-temperature environment, cooling water can be introduced into an annular cooling channel in the heat shield through a water inlet pipeline, the cooling water can flow out through a gap between a water outlet pipeline and an inflation pipeline so as to absorb and take away heat of the high-temperature environment, so that the temperature in the heat shield is kept lower, the electrical lead can be effectively protected from high-temperature damage, the temperature in the inflation pipeline is kept low, the lead of the electrical lead led out from the inflation pipeline is protected from high-temperature damage, meanwhile, low-temperature high-pressure gas can be filled into the heat shield through the inflation pipeline, so that a certain pressure is kept in the heat shield, the gas in the high-temperature environment is prevented from directly entering the heat shield to impact the electric leading device, and the electric leading device is prevented from being damaged.
Drawings
FIG. 1 is a schematic illustration of a last stage low pressure turbine rotor blade strain gage lead out configuration 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 gage lead out structure provided by an embodiment of the present application;
wherein:
1-low pressure turbine outer casing; 2-final stage low pressure turbine rotor disk; 3-last stage low pressure turbine rotor blades; 4-low pressure turbine inner casing; 5-low pressure turbine hollow rotating shaft; 6-rear end support bearing; 7-bearing mounting seats; 8-axis oil collecting ring; 9-thrust augmentation diffuser outer casing; 10-an inner cone; 11-a strain gauge; 12-an electrical starter; 13-a support plate; 14-a heat shield; 15-a water inlet pipe; 16-an outlet conduit; 17-an inflation conduit; 18-water inlet hose; 19-water outlet hose; 20-an insulated duct; 21-a gas supply pipeline; 22-an annular cover; 23-quill.
For a better understanding of the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced, and do not represent actual product dimensions, and 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 solutions and advantages of the present application clearer, the technical solutions of the present application will be further clearly and completely described in the following detailed description with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are only used for explaining the present application, but not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, other related parts may refer to general designs, and the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment without conflict.
In addition, unless otherwise defined, technical or scientific terms used in the description of the present application shall have the ordinary meaning as understood by one of ordinary skill in the art to which the present application belongs. The terms "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", and the like used in the description of the present application, which indicate orientations, are used only to indicate 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 when the absolute position of the object to be described is changed, the relative positional relationships may be changed accordingly, and thus, should not be construed as limiting the present application. The use of "first," "second," "third," and the like in the description of the present application is for descriptive purposes only to distinguish between different components and is not to be construed as indicating or implying relative importance. The use of the terms "a," "an," or "the" and similar referents in the description of the application should not be construed as an absolute limitation of quantity, but rather as the presence of at least one. The word "comprising" or "comprises", and the like, when used in this description, is intended to specify the presence of stated elements or items, but not the exclusion of other elements or items.
Further, it is noted that, unless expressly stated or limited otherwise, the terms "mounted," "connected," and the like are used in the description of the invention in a generic sense, e.g., connected as either a fixed connection or a removable connection or integrally connected; can be mechanically or electrically connected; they may be directly connected or indirectly connected through an intermediate medium, or they may be connected through the inside of two elements, and those skilled in the art can understand their specific meaning in this application according to the specific situation.
The present application is described in further detail below with reference to fig. 1-2.
A whole-aircraft engine low-turbine rotor blade stress measurement structure comprises:
a support plate 13 arranged in the inner cone 10, connected to the bearing mounting seat 7, and having a through hole thereon;
the electricity leading device 12 is arranged in the inner cone 10, the stator part of the electricity leading device is connected to the supporting plate 13, and the rotor part of the electricity leading device passes through the through hole to be connected with the axis oil collecting ring 8 and is connected to a lead wire of the strain gauge 11;
the heat shield 14 is arranged in the inner cone 10, connected to the support plate 13 and used for shielding the electric leading device 12, an annular cooling channel is arranged in the heat shield, the outer wall of the heat shield is provided with a heat shield water inlet hole and a heat shield water outlet hole which are 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 penetrates through water inlet holes in the thrust augmentation diffuser outer casing 9 and the inner cone 10 and is communicated with a water inlet hole of the heat shield;
the inlet end of the water outlet pipeline 16 penetrates through water outlets on the diffuser outer casing 9 and the inner cone 10 and is communicated with the water outlet of the heat insulation cover;
and the outlet end of the air charging pipeline 17 penetrates through the water outlet pipeline 16, the water outlet hole of the heat shield and the annular cooling channel, is communicated with the lead hole of the heat shield, and leads of the electric lead 12 are led out of the air charging pipeline.
When the aircraft engine works, a lead of an electrical lead 12 is led out by using the stress measurement structure of the whole low-pressure turbine rotor blade of the aircraft engine disclosed by the embodiment, so as to realize the acquisition of the stress level and the distribution measurement signal of the final-stage low-pressure turbine rotor blade 3, wherein the electrical lead 12 is covered by the heat insulation cover 14, so that the electrical lead 12 is protected in a high-temperature environment, cooling water can be introduced into an annular cooling channel in the heat insulation cover 14 through the water inlet pipeline 15, the cooling water can flow out through a 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 temperature in the heat insulation cover 14 is kept to be lower, the electrical lead 12 can be effectively protected from high-temperature damage, the air charging pipeline 17 is kept at a low temperature, and the lead of the electrical lead 12 led out in the air charging pipeline 17 is protected from high-temperature damage, meanwhile, low-temperature high-pressure gas can be filled into the heat insulation cover 14 through the inflation pipeline 17, so that a certain pressure is maintained in the heat insulation cover 14, and the gas in a high-temperature environment is prevented from directly entering the heat insulation cover 14 to impact the electric initiator 12 and damage the electric initiator 12.
The stress measurement structure of the whole low-pressure turbine rotor blade of the aircraft engine disclosed by the embodiment can also be applied to leading out of a temperature and stress measurement sensor lead of a final stage low-pressure turbine rotating part of the aircraft engine in an analogy manner.
In some optional embodiments, in the above structure for measuring stress of a whole low-turbine rotor blade of an aircraft engine, 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 insulation cover, and is flexibly connected between the outlet end of the water inlet pipeline 15 and the water inlet hole of the heat insulation cover so as to coordinate the deformation among the outer casing 9 of the booster diffuser, the inner cone 10 and the heat insulation cover 14 and avoid generating larger local stress;
and the water outlet hose 19 is connected between the inlet end of the water outlet pipeline 16 and the water outlet of the heat shield, and is flexibly connected between the inlet end of the water outlet pipeline 16 and the water outlet 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 booster diffuser and avoid generating larger local stress.
In some alternative embodiments, in the above-mentioned stress measurement structure for the whole aircraft engine low-turbine rotor blade, the water inlet pipe 15 and its corresponding booster diffuser outer casing 9 and the water inlet holes on the inner cone 10, the heat shield water inlet holes, and the water inlet hoses 18 are multiple and distributed circumferentially along the heat shield 14.
In some optional embodiments, in the above structure for measuring stress of a whole low-turbine rotor blade of an aircraft engine, further includes:
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 filled with heat insulation materials between the air inflation pipeline 17 and the heat insulation pipeline.
For the aero-engine complete machine low-vortex rotor blade stress measurement structure 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 gap between the water outlet pipe 16 and the heat insulation pipe 20, the heat insulation material is filled between the heat insulation pipe 20 and the air charging pipe 17, so that heat transfer to the air charging pipe 17 can be avoided, and the air charging pipe 17 can be kept at a lower temperature under the condition that air is introduced into the air charging pipe 17.
In some optional embodiments, in the above-mentioned stress measurement structure for the whole low-turbine rotor blade of the aircraft engine, the bearing mount 7 has a cooling air inlet hole, and the cooling air inlet hole is communicated with the space among the support plate 13, the heat shield 14 and the inner cone 10;
the inner cone 10 is provided with a cooling air outlet which is communicated with the space among the support plate, the heat shield 14 and the inner cone 10.
For the aircraft engine complete machine low-turbine rotor blade stress measurement structure disclosed in the above embodiment, it can be understood by those skilled in the art that the cold air in the low-pressure turbine inner casing 4 can enter the space between the heat shield 14 and the inner cone 10 through the bearing mounting seat cooling air inlet hole, and absorb and take away the heat in the space, thereby further ensuring that the temperature in the heat shield 14 can be kept at a lower temperature, and protecting the electrical lead 12 from high-temperature damage.
In some optional embodiments, in the above structure for measuring stress of a whole low-turbine rotor blade of an aircraft engine, further includes:
the outlet end of the air supply pipeline 21 penetrates through air supply holes in 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-stage low-pressure turbine rotor disc 2;
the annular cover 22 is sleeved on the periphery of the axis oil receiving ring 8, is connected with the supporting plate 13 and forms an air supply channel with the supporting plate 13; the ventilation channel is communicated with the space between the bearing mounting seat 7 and the final-stage low-pressure turbine rotor disc 2 through an air supply hole in the bearing mounting seat 7 and is communicated with the inside of the axis oil-collecting ring 8 through an air outlet hole in the axis oil-collecting ring 8.
For the aircraft engine complete machine low-turbine rotor blade stress measurement structure disclosed in the above embodiments, it can be understood by those skilled in the art that, cooling air can be introduced into the space between the bearing mounting seat 7 and the final-stage low-pressure turbine rotor disc 2 through an air supply pipeline 21, the cooling air can enter the air supply channel formed between the annular cover 22 and the support plate 13 through the air supply holes on the bearing mounting seat 7, then the air outlet on the axle center oil collecting ring 8 flows out from the axle center oil collecting ring 8 and the low-pressure turbine hollow rotating shaft 5 to absorb and take away the heat in the air supply channel, the rotor parts of the electrical starter 12 are cooled, the rotor parts of the electrical starter 12 are prevented from being damaged by high temperature, and can maintain a certain pressure in the gas supply channel to prevent gas in a high-temperature environment from entering and impacting the rotor part of the electrical starter 12, the rotor component is damaged, and meanwhile, the flow path can also realize the sealing function of the lubricating oil cavity.
In some optional embodiments, in the above structure for measuring stress of a whole low-turbine rotor blade of an aircraft engine, further includes:
the sleeve shaft 23 is connected between the rotor part of the electrical starter 12 and the axis oil collecting ring 8, a lead of the hollow supply transformer 11 penetrates through the sleeve shaft, effective torque transmission is carried out between the electrical starter 12 and the low-pressure turbine hollow rotating shaft 5, circular arc structures are designed at two ends of the sleeve shaft, the sleeve shaft can be used for compensating different axial degrees between the electrical starter 12 and the low-pressure turbine hollow rotating shaft 5, measuring accuracy is guaranteed, and in addition, the sleeve shaft 23 can be broken and disconnected in time when in failure, and the electrical starter 12 is protected from mechanical damage.
The embodiments are described in a progressive manner in the specification, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other.
Having thus described the present application in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present application is not limited to those specific embodiments, and that equivalent modifications or substitutions of related technical features may be made by those skilled in the art without departing from the principle of the present application, and those modifications or substitutions will fall within the scope of the present application.
Claims (7)
1. The utility model provides an aeroengine whole machine low pressure turbine rotor blade stress measurement structure for draw forth the lead wire of last stage low pressure turbine rotor blade stress measurement strainometer, its characterized in that includes:
the supporting plate (13) is arranged in the inner cone (10), is connected to the bearing mounting seat (7) and is provided with a through hole;
the electricity leading device (12) is arranged in the inner cone (10), the stator part of the electricity leading device is connected to the supporting plate (13), the rotor part of the electricity leading device passes through the through hole to be connected with the axis oil collecting ring (8), and a lead wire connected to the strain gauge (11);
the heat insulation cover (14) is arranged in the inner cone (10), is connected to the supporting plate (13), covers the electric initiator (12), is internally provided with an annular cooling channel, is provided with a heat insulation cover water inlet hole and a heat insulation cover water outlet hole which are communicated with the annular cooling channel on the outer wall, and is provided with a heat insulation cover lead hole communicated with the annular cooling channel on the inner wall;
the outlet end of the water inlet pipeline (15) penetrates through water inlet holes in the thrust augmentation diffuser outer casing (9) and the inner cone (10) and is communicated with the heat insulation cover water inlet hole;
the inlet end of the water outlet pipeline (16) penetrates through water outlet holes in the outer diffuser casing (9) and the inner cone (10) and is communicated with the water outlet hole of the heat insulation cover;
and the outlet end of the air charging pipeline (17) penetrates through the water outlet pipeline (16), the water outlet hole of the heat shield and the annular cooling channel, is communicated with the lead hole of the heat shield, and leads of the electric power generator (12) are led out from the air charging pipeline.
2. The aeroengine whole low turbine rotor blade stress-measuring structure of claim 1,
further comprising:
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 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 insulation cover.
3. The aeroengine whole low turbine rotor blade stress-measuring structure of claim 1,
the water inlet pipelines (15) and the corresponding stressing diffuser outer casing (9) and the water inlet holes, the heat insulation cover water inlet holes and the water inlet hoses (18) on the inner cone (10) are multiple and are distributed along the circumferential direction of the heat insulation cover (14).
4. The aeroengine whole low turbine rotor blade stress-measuring structure of claim 1,
further comprising:
and the heat insulation pipeline (20) is arranged in the water outlet pipeline (16), is sleeved on the periphery of the air inflation pipeline (17), and is filled with heat insulation materials between the air inflation pipeline (17).
5. The aeroengine whole low turbine rotor blade stress-measuring structure of claim 1,
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 support plate (13), the heat insulation cover (14) and the inner cone (10);
the inner cone (10) is provided with a cooling air outlet which is communicated with the space among the support plate (13), the heat shield (14) and the inner cone (10).
6. The aeroengine whole low turbine rotor blade stress-measuring structure of claim 1,
further comprising:
the outlet end of the gas supply pipeline (21) penetrates through gas supply holes in the outer casing (1) and the inner casing (4) of the low-pressure turbine and is communicated with the space between the bearing mounting seat (7) and the final-stage low-pressure turbine rotor disc (2);
the annular cover (22) is sleeved on the periphery of the axis oil collecting ring (8), is connected with the supporting plate (13), and forms an air supply channel with the supporting plate (13); the ventilation channel is communicated with the space between the bearing mounting seat (7) and the final-stage low-pressure turbine rotor disc (2) through an air supply hole in the bearing mounting seat (7), and is communicated with the inside of the axis oil collecting ring (8) through an air outlet hole in the axis oil collecting ring (8).
7. The aeroengine whole low turbine rotor blade stress-measuring structure of claim 1,
further comprising:
and the sleeve shaft (23) is connected between the rotor part of the electrical starter (12) and the axis oil collecting ring (8), and a lead of the supply variable meter (11) passes through the sleeve shaft.
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 |
|---|---|---|---|
| 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 true CN114964787A (en) | 2022-08-30 |
| CN114964787B CN114964787B (en) | 2023-09-22 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115342774A (en) * | 2022-10-20 | 2022-11-15 | 北京航天动力研究所 | Strain measurement system for high-speed flexible rotor turbine disc of liquid rocket engine |
Citations (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4393697A (en) * | 1980-01-31 | 1983-07-19 | Hitachi, Ltd. | Air flow rate measuring apparatus |
| GB9014619D0 (en) * | 1990-06-30 | 1990-08-22 | Rolls Royce Plc | Dual readout temperature and strain sensor combinations |
| CN106441905A (en) * | 2016-08-16 | 2017-02-22 | 中国航空工业集团公司沈阳发动机设计研究所 | Signal transmission device with rotor acting as benchmark |
| WO2019005467A2 (en) * | 2017-06-09 | 2019-01-03 | Zornes David Allen | Tip-path airfoil thrust production in rotary-wing aircraft |
| CN110261117A (en) * | 2018-03-12 | 2019-09-20 | 中国航发商用航空发动机有限责任公司 | Turbogenerator simulation experiment system |
| CN111337262A (en) * | 2020-04-16 | 2020-06-26 | 中国航发沈阳发动机研究所 | Engine rotor test lead structure |
| CN113565583A (en) * | 2021-07-19 | 2021-10-29 | 中国航发沈阳发动机研究所 | Device for testing dynamic stress of complete high-pressure turbine rotor of double-rotor turbofan engine |
| CN113654701A (en) * | 2021-08-24 | 2021-11-16 | 中国航发湖南动力机械研究所 | Dynamic stress measuring device for rotor blade of aircraft engine and application of dynamic stress measuring device |
| CN113844677A (en) * | 2021-10-25 | 2021-12-28 | 中国航发沈阳发动机研究所 | Axial lead structure for measuring dynamic stress of whole high-pressure turbine of turbofan engine |
| CN113984259A (en) * | 2021-09-18 | 2022-01-28 | 中国航发南方工业有限公司 | Lead device for measuring dynamic stress of compressor blade |
| CN114088231A (en) * | 2021-11-05 | 2022-02-25 | 中国航发沈阳发动机研究所 | Temperature testing device for rotor part of low-pressure turbine in complete state of aircraft engine |
| CN114166393A (en) * | 2021-11-10 | 2022-03-11 | 中国航发湖南动力机械研究所 | Blade dynamic stress measuring structure |
| CN114295382A (en) * | 2021-12-22 | 2022-04-08 | 北京动力机械研究所 | Air inlet casing of high-pressure turbine blade vibration stress dynamic measurement core machine tester |
-
2022
- 2022-05-12 CN CN202210518372.8A patent/CN114964787B/en active Active
Patent Citations (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4393697A (en) * | 1980-01-31 | 1983-07-19 | Hitachi, Ltd. | Air flow rate measuring apparatus |
| GB9014619D0 (en) * | 1990-06-30 | 1990-08-22 | Rolls Royce Plc | Dual readout temperature and strain sensor combinations |
| CN106441905A (en) * | 2016-08-16 | 2017-02-22 | 中国航空工业集团公司沈阳发动机设计研究所 | Signal transmission device with rotor acting as benchmark |
| WO2019005467A2 (en) * | 2017-06-09 | 2019-01-03 | Zornes David Allen | Tip-path airfoil thrust production in rotary-wing aircraft |
| CN110261117A (en) * | 2018-03-12 | 2019-09-20 | 中国航发商用航空发动机有限责任公司 | Turbogenerator simulation experiment system |
| CN111337262A (en) * | 2020-04-16 | 2020-06-26 | 中国航发沈阳发动机研究所 | Engine rotor test lead structure |
| CN113565583A (en) * | 2021-07-19 | 2021-10-29 | 中国航发沈阳发动机研究所 | Device for testing dynamic stress of complete high-pressure turbine rotor of double-rotor turbofan engine |
| CN113654701A (en) * | 2021-08-24 | 2021-11-16 | 中国航发湖南动力机械研究所 | Dynamic stress measuring device for rotor blade of aircraft engine and application of dynamic stress measuring device |
| CN113984259A (en) * | 2021-09-18 | 2022-01-28 | 中国航发南方工业有限公司 | Lead device for measuring dynamic stress of compressor blade |
| CN113844677A (en) * | 2021-10-25 | 2021-12-28 | 中国航发沈阳发动机研究所 | Axial lead structure for measuring dynamic stress of whole high-pressure turbine of turbofan engine |
| CN114088231A (en) * | 2021-11-05 | 2022-02-25 | 中国航发沈阳发动机研究所 | Temperature testing device for rotor part of low-pressure turbine in complete state of aircraft engine |
| CN114166393A (en) * | 2021-11-10 | 2022-03-11 | 中国航发湖南动力机械研究所 | Blade dynamic stress measuring structure |
| CN114295382A (en) * | 2021-12-22 | 2022-04-08 | 北京动力机械研究所 | Air inlet casing of high-pressure turbine blade vibration stress dynamic measurement core machine tester |
Non-Patent Citations (2)
| Title |
|---|
| 杨春华等: "某型发动机涡轮转子温度和应力测试研究", 《航空发动机》, vol. 43, no. 2 * |
| 贾瑞琦等: "航空发动机整机静子建模与刚度验证", 《航空发动机》, vol. 41, no. 6 * |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115342774A (en) * | 2022-10-20 | 2022-11-15 | 北京航天动力研究所 | Strain measurement system for high-speed flexible rotor turbine disc of liquid rocket engine |
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