CN117890117A - Aeroengine rear end low pressure rotor test system - Google Patents
Aeroengine rear end low pressure rotor test system Download PDFInfo
- Publication number
- CN117890117A CN117890117A CN202211223703.1A CN202211223703A CN117890117A CN 117890117 A CN117890117 A CN 117890117A CN 202211223703 A CN202211223703 A CN 202211223703A CN 117890117 A CN117890117 A CN 117890117A
- Authority
- CN
- China
- Prior art keywords
- low pressure
- low
- pressure shaft
- rear end
- shaft
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 238000012360 testing method Methods 0.000 title claims abstract description 36
- 230000008054 signal transmission Effects 0.000 claims abstract description 43
- WABPQHHGFIMREM-UHFFFAOYSA-N lead(0) Chemical compound [Pb] WABPQHHGFIMREM-UHFFFAOYSA-N 0.000 claims abstract description 37
- 230000006698 induction Effects 0.000 claims abstract description 8
- 238000007789 sealing Methods 0.000 claims description 15
- 238000004891 communication Methods 0.000 claims description 2
- 238000007599 discharging Methods 0.000 claims description 2
- 238000005516 engineering process Methods 0.000 abstract description 2
- 238000012857 repacking Methods 0.000 description 4
- 244000126211 Hericium coralloides Species 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 230000000903 blocking effect Effects 0.000 description 2
- 238000002485 combustion reaction Methods 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 238000005259 measurement Methods 0.000 description 2
- 230000002093 peripheral effect Effects 0.000 description 2
- 238000012545 processing Methods 0.000 description 2
- 230000004888 barrier function Effects 0.000 description 1
- 230000005540 biological transmission Effects 0.000 description 1
- 239000004020 conductor Substances 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 230000005611 electricity Effects 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 238000001746 injection moulding Methods 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 238000010030 laminating Methods 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 230000000149 penetrating effect Effects 0.000 description 1
- 239000011148 porous material Substances 0.000 description 1
- 238000009420 retrofitting Methods 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
Landscapes
- Measuring Fluid Pressure (AREA)
Abstract
The invention discloses a rear end low-voltage rotor testing system of an aero-engine, which relates to the field of aero-engine testing technology and comprises the following components: an aero-engine; the testing equipment comprises an induction part, a lead wire and a signal transmission device; the induction part is fixed on a low-pressure rotor of the aero-engine; the signal transmission device is arranged at the rear end of the low-pressure shaft of the aero-engine; the two ends of the lead are respectively connected with the induction part and the signal transmission device; the signal transmission device is arranged at the rear end of the low-voltage shaft, the distance between the signal transmission device and the sensing part is reduced, and the length of the corresponding lead wire for connection is reduced, so that the lead wire is not easy to damage in a working environment due to overlong length, and meanwhile, the loss of the signal when transmitted from the sensing part to the signal transmission device is reduced.
Description
Technical Field
The invention relates to the field of aero-engine testing technology, in particular to a rear-end low-pressure rotor testing system of an aero-engine.
Background
In the working process of the aeroengine, the rotor bears a high-temperature and high-pressure environment and rotates at a high speed, and the accurate measurement of performance data such as vibration stress, temperature and the like of the rotor is not only the requirements of improving the development level and enhancing the performance and reliability of the engine, but also an important assessment item of the airworthiness standard.
In general, the complete machine test of the aero-engine is limited by the space for structural modification, and the rotor measurement can only be aimed at low-voltage components. The measuring device comprises a sensing part arranged on the rotor of the low-pressure shaft and a signal transmission device arranged at the front end of the low-pressure shaft (namely, the side of the low-pressure shaft close to the fan). The scheme is easy to realize for refitting the engine, but when being used for testing the low-pressure turbine component positioned at the rear side of the low-pressure shaft, the signal transmission device is arranged at the front end of the engine, so that the lead wire length is too long, the engine is easy to damage in the working environment, and the accuracy of a test signal is affected.
Disclosure of Invention
The invention aims to overcome the defects that in the prior art, when low-pressure turbine performance data are measured, the length of a lead wire between a sensing part and a signal transmission device is too long, the lead wire is easy to damage in a working environment, the accuracy of a test signal is affected and the like, and provides a rear-end low-pressure rotor test system of an aeroengine.
The invention solves the technical problems by the following technical scheme:
the invention provides an aero-engine rear end low-pressure rotor test system, which comprises:
an aero-engine;
the testing equipment comprises an induction part, a lead wire and a signal transmission device;
the induction part is fixed on a low-pressure rotor of the aero-engine;
the signal transmission device is arranged at the rear end of the low-pressure shaft of the aero-engine;
and two ends of the lead are respectively connected with the sensing part and the signal transmission device.
In the scheme, the sensing part is fixed on the low-voltage rotor and used for obtaining parameters of the low-voltage rotor, and the signal transmission device is electrically connected with the sensing part through a lead wire so as to transmit the parameters obtained by the sensing part to the outside of the aeroengine for calculation and processing; compared with the mode that the signal transmission device is arranged at the front end of the aeroengine, the signal transmission device is arranged at the rear end of the low-voltage shaft, the distance between the signal transmission device and the sensing part is reduced, and the length of a corresponding lead wire used for connection is reduced, so that the lead wire is not easy to damage in a working environment due to overlong length, and meanwhile, the loss of signals transmitted from the sensing part to the signal transmission device is reduced.
Preferably, the signal transmission device comprises a rotating part and a stationary part; the rotating part is fixed at the rear end of the low-pressure shaft of the aero-engine; the stationary part is fixed on a stator assembly at the rear end of the aeroengine; the stationary portion is electrically connected to the rotating portion.
In the scheme, the rotating part of the signal transmission device is fixed at the rear end of the low-voltage shaft and is electrically connected with the induction part through a wire, and the fixed part can rotate together with the low-voltage shaft and the rotor, so that the lead is not easy to wind and tie; the static part is fixed on the stator assembly so as to transmit the received signal to the outside of the aeroengine; the rotating part can rotate relative to the stationary part, and meanwhile the rotating part is kept electrically connected with the stationary part, so that the signal is transmitted from the sensing part to the stationary part.
Preferably, a rotary joint is arranged at the rear end of the low-pressure shaft, and the rotary part is inserted into the rotary joint and fixedly connected with the low-pressure shaft.
In the scheme, the rotating part is inserted into the rotating port of the low-pressure shaft to realize the fixed connection between the rotating part and the rear end of the low-pressure shaft.
Preferably, a lead hole is formed in the circumferential side of the low-voltage shaft, one end of the lead is fixed to the sensing portion, the other end of the lead extends along the outer circumferential surface of the low-voltage shaft, enters the low-voltage shaft through the lead hole and passes through the rotating port to be connected with the rotating portion, and the lead is connected with the lead hole in an airtight manner.
In the scheme, a lead hole is formed in the peripheral side of the low-voltage shaft, and a lead enters the inner side of the low-voltage shaft from the outer side of the low-voltage shaft through the lead hole and passes through the interface and then is connected with the rotating part; the connection between the low-voltage shaft and the signal transmission device is realized through the connection between the low-voltage shaft and the signal transmission device, so that the connection between the lead and the rotating part is facilitated, and the probability of damage of the lead when the lead rotates along with the low-voltage shaft and the low-voltage rotor is reduced; the lead and the hole wall of the lead hole are sealed, so that the sealing condition of the inner side and the outer side of the low-voltage shaft is not easy to be damaged when the lead hole is formed.
Preferably, the lead wire is bonded and fixed to the rear surface of the low-voltage rotor.
In this scheme, the lead wire laminating is fixed at the rear surface of low pressure rotor for the lead wire is comparatively stable when rotating along with low pressure rotor, makes the lead wire not fragile.
Preferably, the aero-engine rear end low-pressure rotor test system further comprises a refitted pipeline; the periphery of the rear end of the low pressure shaft is provided with a vent hole, one end of the refitted pipeline is connected with the vent hole in an airtight mode, and the refitted pipeline is used for discharging air flow in the low pressure shaft.
In the scheme, the refitted pipeline is communicated with the vent hole of the low-pressure shaft, and air flow in the low-pressure shaft can flow into the refitted pipeline through the vent hole and be discharged out of the aeroengine through the refitted pipeline, so that the normal pneumatic condition of the air path of the axis of the low-pressure shaft is ensured, and the requirement of sealing pressure difference is not influenced.
Preferably, a wiring hole is formed in the rotating portion, which is inserted into one end of the switching port, and the lead is inserted into the wiring hole in an airtight manner.
In this scheme, the lead wire can insert in the wiring hole after wearing out through the switching mouth and realize being connected with the electricity of rotating part, and the lead wire is inserted in the wiring hole simultaneously with the gas tightness for the air current of the axle center of low pressure axle is difficult for leaking from the switching mouth, thereby further makes the low pressure axle satisfy the demand of sealing pressure differential.
Preferably, an ejector for exhausting air in the refitted pipeline is arranged at one end of the refitted pipeline far away from the low-pressure shaft.
In this scheme, through setting up the ejector in the one end that the low pressure axle was kept away from to the repacking pipeline to be convenient for discharge the air current in the repacking pipeline outside the aeroengine.
Preferably, the retrofit pipeline comprises a bypass pipe and a retrofit connector, the low pressure shaft is rotatably arranged in the retrofit connector, and the vent hole is communicated with the inner cavity of the retrofit connector; the bypass pipe is fixed on a stator assembly of the aeroengine and is communicated with the inner cavity of the refitted joint and the external atmosphere.
In this solution, the retrofit pipeline comprises a bypass pipe and a retrofit connector, the retrofit connector and the bypass pipe being fixed relative to the stator assembly, such that the end of the low pressure shaft inserted into the retrofit connector rotates relative to the retrofit connector; the inner cavity of the low-pressure shaft is communicated with the inner cavity of the refitting joint through the lead hole, and the bypass pipe is communicated with the inner cavity of the refitting joint, so that air flow in the low-pressure shaft can be continuously discharged into the external atmosphere through the refitting joint and the bypass pipe.
Preferably, the periphery of the rear end of the low pressure shaft is provided with a sealing comb tooth, and the sealing comb tooth is used for sealing the periphery of the low pressure shaft and the inner cavity of the refitting joint.
In this scheme, after the rear end of low pressure axle inserts in the repacking joint, through the sealing comb tooth of low pressure axle rear end seal the rear end of low pressure axle with the connector of repacking joint, guarantee the seal between aeroengine's inner chamber and the low pressure axle.
Preferably, the lead hole is located on a side of the retrofit connector adjacent to the low pressure rotor.
In the scheme, the lead hole is formed in one side, close to the low-voltage rotor, of the modified connector, so that the position of the lead hole is staggered from the position of the low-voltage shaft inserted into the modified connector, and interference between the lead hole and the modified connector during lead arrangement is avoided.
The invention has the positive progress effects that:
compared with the mode that the signal transmission device is arranged at the front end of the aeroengine, the low-voltage rotor testing system for the rear end of the aeroengine is characterized in that the signal transmission device is arranged at the rear end of the low-voltage shaft, the distance between the signal transmission device and the sensing part is reduced, and the length of a corresponding lead wire used for connection is reduced, so that the lead wire is not easy to damage in a working environment due to overlong length, and meanwhile loss when signals are transmitted from the sensing part to the signal transmission device is reduced.
Drawings
Fig. 1 is a schematic structural view of an aeroengine according to an embodiment of the present invention.
Fig. 2 is a schematic structural view of the aft end of an aircraft engine according to an embodiment of the present invention.
Fig. 3 is a schematic structural view of the rear end of the low pressure shaft according to an embodiment of the present invention.
FIG. 4 is a schematic cross-sectional view of the rear end of a low pressure shaft according to an embodiment of the present invention.
Reference numerals illustrate:
aeroengine 100
Casing 110
Fan 120
Low pressure compressor 130
High pressure compressor 140
Combustion chamber 150
High pressure turbine 160
Low pressure turbine 170
Low pressure shaft rear bearing 171
Tail spray assembly 180
Stator assembly 181
Low pressure rear bearing support 182
Low pressure shaft 190
Interface 191
Barrier plate 192
Lead hole 193
Vent 194
Sealing comb 195
Low pressure rear bearing chamber 101
Sensing part 210
Signal transmission device 220
Rotating portion 221
Stationary portion 222
Lead 230
Retrofit pipeline 300
Retrofit connector 310
Bypass pipe 320
Tail end exhaust pipe 330
Ejector 340
Jet nozzle 341
Tapered tube section 342
Detailed Description
The invention is further illustrated by means of examples which follow, without thereby restricting the scope of the invention thereto.
The embodiment discloses an aero-engine rear end low-pressure rotor test system, referring to fig. 1 and 2, the aero-engine rear end low-pressure rotor test system comprises an aero-engine 100 and test equipment arranged in the aero-engine 100, wherein the test equipment is used for measuring data parameters of a low-pressure rotor of the aero-engine 100 and transmitting the data parameters to the outside of the aero-engine 100.
Referring to fig. 1, an aircraft engine 100 includes a casing 110 and a fan 120, a low-pressure compressor 130, a high-pressure compressor 140, a combustion chamber 150, a high-pressure turbine 160, a low-pressure turbine 170, and a tail jet assembly 180 disposed within the casing 110 and disposed in sequence from a front end to a rear end of the aircraft engine 100. The center of the aeroengine 100 is provided with a low pressure shaft 190 and a high pressure shaft sleeved outside the low pressure shaft 190. Wherein the rotor of the fan 120, the rotor of the low pressure compressor 130, and the rotor of the low pressure turbine 170 are all fixed on the low pressure shaft 190, the rotor of the fan 120, the rotor of the low pressure compressor 130, and the rotor of the low pressure turbine 170 are collectively referred to as a low pressure rotor.
Wherein, tail spout assembly 180 internal fixation is provided with stator subassembly 181. The stator assembly 181 includes a low pressure rear bearing support 182, a low pressure shaft rear bearing 171 is provided on the low pressure turbine 170, and the low pressure rear bearing support 182 is connected to the low pressure shaft rear bearing 171, so that a low pressure rear bearing chamber 101 is formed between the low pressure rear bearing support 182 and the low pressure shaft 190 and between the low pressure turbine 170, thereby enabling rotational contact between the stator assembly 181 and the low pressure turbine 170.
The front end of the aero-engine 100 refers to a portion into which the airflow flows, and the rear end of the aero-engine 100 refers to a portion from which the airflow flows out of the aero-engine 100.
The aft end low pressure rotor test system of the present embodiment is used to test parameters related to the low pressure rotor of the aft end of the aircraft engine 100, i.e., parameters related to the rotor of the low pressure turbine 170.
Referring to fig. 2, the test equipment includes a sensing portion 210 for acquiring parameters of a low pressure rotor at the aft end of the aircraft engine 100 and a signal transmission device 220 for receiving the parameters of the sensing portion 210 and transmitting the data to other equipment outside the aircraft engine 100 for processing. The sensing part 210 is fixedly provided on the rotor of the low pressure turbine 170, and the signal transmission device 220 is provided at the rear end of the low pressure shaft 190. A lead 230 is connected between the sensing part 210 and the signal transmission device 220 so as to transmit the data obtained by the sensing part 210 to the signal transmission device 220.
Thereby, the distance between the signal transmission device 220 and the sensing part 210 is reduced, and the length of the lead wire 230 for connecting the sensing part 210 and the signal transmission device 220 is reduced, so that the lead wire 230 is not easily damaged in the working environment due to the overlong length, and the loss when the signal is transmitted from the sensing part 210 to the signal transmission device 220 is reduced.
The rear end of the low pressure shaft 190 refers to a portion of the low pressure shaft 190 near the rear end of the aero-engine 100, and in this embodiment, the rear end of the low pressure shaft 190 is an axial exhaust pipe.
The sensing portion 210 may be a temperature sensor, a pressure sensor, a strain sensor, etc., and may be specifically selected according to the tested parameters.
Referring to fig. 2, the signal transmission device 220 includes a rotating portion 221 and a stationary portion 222. The stationary part 222 is connected to the stator assembly 181, and in particular, the stationary part 222 is fixed to the low-pressure rear bearing support 182 by bolts in this embodiment. The rotating portion 221 is fixedly coupled to the rear end of the low voltage shaft 190 and is coupled to the lead wire 230 of the connection sensing part 210.
The rotating portion 221 and the stationary portion 222 are electrically connected, so that data acquired by the sensing unit 210 can be transmitted to the rotating portion 221 via the lead wire 230, then transmitted to the stationary portion 222, and transmitted to the outside of the aeroengine 100 via the stationary portion 222. In this embodiment, the rotating part 221 and the stationary part 222 are both made of conductors, the rotating part 221 is cylindrical, the stationary part 222 is cylindrical, and one end of the rotating part 221, which is far away from the low-pressure shaft 190, is inserted into the stationary part 222, and the rotating part 221 can rotate relative to the stationary part 222 and maintain sliding contact with the stationary part 222, so that the rotating part 221 can be conducted with the stationary part 222, and transmission of an electrical signal is achieved in a physical contact manner. In other embodiments, the rotary part 221 and the stationary part 222 may also transmit electric signals by means of telemetry.
Wherein, the rotating portion 221 is provided with a wiring hole, and the lead 230 can be inserted into the rotating portion 221 through the wiring hole to electrically connect the lead 230 with the rotating portion 221. Preferably, the air-tightness between the lead wire 230 and the wire connection hole is provided so that the air flow in the axial direction of the low-pressure shaft 190 is not easily entered into the rotating portion 221.
In this embodiment, the current that powers the sensing portion 210 is transferred to the rotating portion 221 via the stationary portion 222 and to the sensing portion 210 via the rotating portion 221. In other embodiments, a power source such as a battery may be provided on the rotor of the low pressure turbine 170 to power the sensing portion 210.
Referring to fig. 2 and 3, the inside of the low pressure shaft 190 is hollow to allow the air flow. The rear end of the low pressure shaft 190 is provided with a rotary joint 191, and one end of the rotary part 221 is inserted into the rotary joint 191 to be fixedly connected with the low pressure shaft 190.
The rotating portion 221 and the low pressure shaft 190 may be connected by pin connection, screw connection, key connection, coupling connection, interference fit connection, riveting, welding, etc. to be suitable for mounting the rotating portion 221 of the signal transmission device 220 of different types. In this embodiment, the shape of the rotating portion 221 is the same as the shape of the adapting opening 191, and the rotating portion 221 and the adapting opening 191 are in interference fit, so that the rotating portion 221 and the adapting opening 191 are sealed while the rotating portion 221 is fixedly connected with the low pressure shaft 190, and the sealing pressure difference at the adapting opening 191 of the low pressure shaft 190 meets the design requirement. In addition, in other embodiments, the rotary part 221 and the adapter 191 may be sealed by other manners.
Referring to fig. 2 to 4, an annular blocking plate 192 is provided at an inner side of a rear end of the low pressure shaft 190 such that, after the rotating portion 221 is inserted into the swivel joint 191, the blocking plate 192 abuts against an end surface of an end of the rotating portion 221 inserted into the low pressure shaft 190, thereby facilitating positioning and placement of the rotating portion 221.
Wherein a plurality of lead holes 193 are opened at the outer circumference of the rear end of the low pressure shaft 190. The lead hole 193 communicates the inner cavity of the low pressure shaft 190 with the outside of the low pressure shaft 190. The connection between the lead wire 230 and the sensing portion 210 is located at the outer side of the low voltage shaft 190, the lead wire 230 can penetrate into the low voltage shaft 190 through the lead wire hole 193, and one end of the lead wire 230, which is far away from the sensing portion 210, is connected to the rotating portion 221 through the swivel joint 191. Thereby achieving electrical connection of the sensing part 210 to the rotating part 221 while facilitating connection of the lead wire 230 to the rotating part 221.
The shape and number of the lead holes 193 may be designed such that the total effective flow area thereof is not changed, for example, when the lead holes 193 are circular through holes, the effective flow area of the lead holes 193 may be calculated by the following formula:
effective flow area = n pi r 2
Where n is the number of holes and r is the diameter of the holes.
The lead 230 is hermetically, i.e., hermetically, connected to the wall of the lead hole 193, so that the opening of the lead hole 193 is unlikely to break the sealing condition of the inner and outer sides of the low pressure shaft 190. Specifically, the seal between the leads 230 and the lead holes 193 may be achieved herein by way of an injection molding seal, an interference fit seal, or the like.
The portion of the lead wire 230 at the rotor of the low pressure turbine 170 is fixedly disposed against the rear side surface of the rotor of the low pressure turbine 170, and the portion of the lead wire 230 outside the low pressure shaft 190 is fixedly disposed against the outer peripheral surface of the low pressure shaft 190. This arrangement enables lead wires 230 to be compactly and securely arranged while lead wires 230 are less susceptible to damage as they rotate with low pressure turbine 170 rotor and low pressure shaft 190. The extending direction of the lead wire 230 on the low voltage shaft 190 is preferably parallel to the axial direction of the low voltage shaft 190, thereby reducing the amount of the lead wire 230
Referring to fig. 2 to 4, the outer circumference of the rear end of the low pressure shaft 190 is provided with a plurality of vent holes 194 communicating the inner cavity of the low pressure shaft 190 with the outside of the low pressure shaft 190, and the plurality of vent holes 194 are arranged at intervals along the circumference of the low pressure shaft 190. The aircraft engine aft end low pressure rotor test system also includes a retrofit line 300, the retrofit line 300 being in communication with the vent 194 of the low pressure shaft 190 and the junction of the retrofit line 300 and the low pressure shaft 190 being sealed. Therefore, the air flow in the low-pressure shaft 190 can flow into the modified pipeline 300 through the air vent 194 and be discharged out of the aeroengine 100 through the modified pipeline 300, so that the normal air passage pneumatic condition of the axis of the low-pressure shaft 190 is ensured, and the requirement of sealing pressure difference is not influenced.
Wherein the retrofit line 300 comprises a retrofit connector 310 and a bypass pipe 320. The front and rear ends of the adapter 310 are each provided with a through hole penetrating the inner cavity thereof, and the rear end of the low pressure shaft 190 is inserted into the adapter 310 from the through hole at the front end of the adapter 310 so that the inner cavity of the adapter 310 communicates with the vent 194. One end of the bypass tube 320 communicates with the interior cavity of the retrofit adapter 310 and the other end communicates with the outside atmosphere. Thus, the airflow in the interior cavity of the low pressure shaft 190 can be vented through the vent 194 into the interior cavity of the retrofit connector 310 and then out through the bypass tube 320 to the outside atmosphere. Thereby reducing the effect of the arrangement of the signal transmission device 220 on the airflow within the low pressure shaft 190.
Wherein the lead hole 193 is located outside of the side of the retrofit connector 310 that is adjacent to the rotor of the low pressure turbine 170, thereby staggering the position of the lead hole 193 from the position of the low pressure shaft 190 inserted into the retrofit connector 310, avoiding interference with the retrofit connector 310 when the lead 230 is deployed.
The periphery of low pressure shaft 190 is provided with the seal comb 195, and low pressure shaft 190 seals the clearance between the perforated pore wall of low pressure shaft 190 and adapter 310 through the seal comb 195 for seal between bearing cavity 101 and the low pressure shaft 190 behind the low pressure, guarantee that the pressure differential of sealing satisfies the design requirement.
In this embodiment, the bypass pipe 320 is U-shaped, so as to reserve space for installation of the signal transmission device 220, and avoid interference between the bypass pipe 320 and the signal transmission device 220. The number of the bypass pipes 320 is two, the U-shaped openings of the two bypass pipes 320 are opposite, and the two bypass pipes 320 are respectively located at the upper side and the lower side of the signal transmission device 220. The two bypass pipes 320 communicate with each other at an end remote from the low pressure shaft 190.
The retrofitting pipeline 300 further includes a tail gas exhaust pipe 330, wherein the tail gas exhaust pipe 330 is communicated with one ends of the two bypass pipes 320 far away from the low pressure shaft 190, and the tail gas exhaust pipe 330 is communicated with the outside atmosphere. Whereby the air flow in the bypass pipe 320 can be discharged outside the aircraft engine 100 via the aft tailpipe 330.
Wherein an ejector 340 is disposed in the tail end exhaust pipe 330. Specifically, the ejector 340 adopts a venturi structure, including a jet nozzle 341 and a tapered pipe section 342. The jet nozzle 341 ejects high-speed air flow toward the tapered pipe section 342, so that the air flows in the two bypass pipes 320 are ejected into the tapered pipe section 342 and discharged out of the tail end exhaust pipe 330.
While specific embodiments of the invention have been described above, it will be appreciated by those skilled in the art that this is by way of example only, and the scope of the invention is defined by the appended claims. Various changes and modifications to these embodiments may be made by those skilled in the art without departing from the principles and spirit of the invention, but such changes and modifications fall within the scope of the invention.
Claims (11)
1. An aircraft engine aft end low pressure rotor test system, comprising:
an aero-engine;
the testing equipment comprises an induction part, a lead wire and a signal transmission device;
the induction part is fixed on a low-pressure rotor of the aero-engine;
the signal transmission device is arranged at the rear end of the low-pressure shaft of the aero-engine;
and two ends of the lead are respectively connected with the sensing part and the signal transmission device.
2. The aircraft engine aft end low pressure rotor testing system of claim 1, wherein said signal transmission means comprises a rotating portion and a stationary portion; the rotating part is fixed at the rear end of the low-pressure shaft of the aero-engine; the stationary part is fixed on a stator assembly at the rear end of the aeroengine; the stationary portion is electrically connected to the rotating portion.
3. The aircraft engine rear end low pressure rotor testing system of claim 2, wherein a swivel port is provided at the rear end of the low pressure shaft, and the swivel portion is inserted into the swivel port and fixedly connected to the low pressure shaft.
4. The test system for a low-pressure rotor at the rear end of an aeroengine as claimed in claim 3, wherein a lead hole is formed in the circumferential side of the low-pressure shaft, one end of the lead wire is fixed to the sensing part, the other end of the lead wire extends along the outer circumferential surface of the low-pressure shaft, enters the low-pressure shaft through the lead hole and passes through the transfer port to be connected with the rotating part, and the lead wire is connected with the lead hole in an airtight manner.
5. The aircraft engine aft end low pressure rotor test system of claim 2, wherein said leads are bonded and secured to an aft surface of said low pressure rotor.
6. The aircraft engine aft end low pressure rotor test system of claim 4, further comprising a retrofit line;
the periphery of the rear end of the low pressure shaft is provided with a vent hole, one end of the refitted pipeline is connected with the vent hole in an airtight mode, and the refitted pipeline is used for discharging air flow in the low pressure shaft.
7. The aircraft engine rear end low pressure rotor test system according to claim 6, wherein a wiring hole is provided in an inside of an end of the rotating portion inserted into the adapter, and the lead wire is air-tightly inserted into the wiring hole.
8. The aircraft engine aft end low pressure rotor test system of claim 6, wherein an end of said retrofit piping remote from said low pressure shaft is provided with an ejector for exhausting air within said retrofit piping.
9. The aircraft engine aft end low pressure rotor testing system of claim 6, wherein said retrofit piping includes a bypass pipe and a retrofit connector, said aft end of said low pressure shaft rotatably disposed within said retrofit connector, and said vent hole in communication with an interior cavity of said retrofit connector; the bypass pipe is fixed on a stator assembly of the aeroengine and is communicated with the inner cavity of the refitted joint and the external atmosphere.
10. The aircraft engine aft end low pressure rotor testing system of claim 9, wherein a sealing comb is provided on an outer periphery of an aft end of the low pressure shaft, the sealing comb for sealing the outer periphery of the low pressure shaft to an inner cavity of the retrofit joint.
11. The aircraft engine aft end low pressure rotor testing system of claim 9, wherein said lead hole is located on a side of said retrofit connector proximate said low pressure rotor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211223703.1A CN117890117A (en) | 2022-10-08 | 2022-10-08 | Aeroengine rear end low pressure rotor test system |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211223703.1A CN117890117A (en) | 2022-10-08 | 2022-10-08 | Aeroengine rear end low pressure rotor test system |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117890117A true CN117890117A (en) | 2024-04-16 |
Family
ID=90641802
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211223703.1A Pending CN117890117A (en) | 2022-10-08 | 2022-10-08 | Aeroengine rear end low pressure rotor test system |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117890117A (en) |
-
2022
- 2022-10-08 CN CN202211223703.1A patent/CN117890117A/en active Pending
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN113654701B (en) | Dynamic stress measuring device for aero-engine rotor blade and application thereof | |
| CN113565583B (en) | Device for testing dynamic stress of complete high-pressure turbine rotor of double-rotor turbofan engine | |
| CN108760187B (en) | Blade cracking state monitoring method and system and blade | |
| CN210426948U (en) | Hypersonic wind tunnel model bottom pressure measuring device | |
| CN111982523B (en) | Exhaust test structure | |
| CN117890117A (en) | Aeroengine rear end low pressure rotor test system | |
| CN114166393B (en) | Blade dynamic stress measuring structure | |
| CN113844677A (en) | Axial lead structure for measuring dynamic stress of whole high-pressure turbine of turbofan engine | |
| CN112345105B (en) | Lead structure for testing temperature of rotor disc body of air compressor | |
| CN115163201A (en) | Device and system for testing high-pressure turbine rotor of aircraft engine | |
| CN114964787B (en) | Aeroengine complete machine low vortex rotor blade stress measurement structure | |
| CN108757465B (en) | Compression cavity dynamic pressure measuring device of rotary vane type automobile air conditioner compressor | |
| US11473480B2 (en) | Instrumented turbine exhaust duct | |
| CN115184017A (en) | Aeroengine high pressure rotor testing arrangement and aeroengine test piece | |
| US11041404B2 (en) | In-situ wireless monitoring of engine bearings | |
| CN115184021A (en) | Device and system for testing low-pressure turbine rotor of aircraft engine | |
| CN110470434A (en) | A kind of air leakage test device and the test method of turbocharger | |
| CN219084331U (en) | Test mechanism for checking tightness of steam turbine valve | |
| CN116669378A (en) | Power primer cooling device for aeroengine dynamic stress measurement test | |
| CN112423554B (en) | Water cooling device for data acquisition sensor of rotary combustion chamber | |
| CN111766027A (en) | Radome sealing inspection device and method | |
| US20120171037A1 (en) | Probe assembly for use in turbine engines and method of assembling same | |
| CN117888999A (en) | Refitting structure of aero-engine rear end system | |
| US11933184B2 (en) | Turbomachines with SAW or BAW devices, measuring arrangements and installation methods | |
| CN214384170U (en) | High-performance and high-precision shell convenient to install for turbocharger |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |