CN117646911A - ECT flame monitoring sensor attached to aeroengine combustion chamber - Google Patents
ECT flame monitoring sensor attached to aeroengine combustion chamber Download PDFInfo
- Publication number
- CN117646911A CN117646911A CN202311701381.1A CN202311701381A CN117646911A CN 117646911 A CN117646911 A CN 117646911A CN 202311701381 A CN202311701381 A CN 202311701381A CN 117646911 A CN117646911 A CN 117646911A
- Authority
- CN
- China
- Prior art keywords
- electrodes
- combustion chamber
- shielding layer
- ect
- wall
- 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
- 238000002485 combustion reaction Methods 0.000 title claims abstract description 95
- 238000012544 monitoring process Methods 0.000 title claims abstract description 27
- 230000005540 biological transmission Effects 0.000 claims abstract description 13
- 230000002093 peripheral effect Effects 0.000 claims abstract description 7
- 238000005259 measurement Methods 0.000 abstract description 10
- 230000009286 beneficial effect Effects 0.000 abstract description 5
- 230000035945 sensitivity Effects 0.000 abstract description 3
- 238000005516 engineering process Methods 0.000 description 13
- 238000001514 detection method Methods 0.000 description 8
- 238000000034 method Methods 0.000 description 8
- 230000000007 visual effect Effects 0.000 description 6
- 239000000446 fuel Substances 0.000 description 5
- 238000004868 gas analysis Methods 0.000 description 5
- 239000002184 metal Substances 0.000 description 5
- 229910052751 metal Inorganic materials 0.000 description 5
- 238000003325 tomography Methods 0.000 description 4
- 238000003384 imaging method Methods 0.000 description 3
- 230000005855 radiation Effects 0.000 description 3
- 238000011160 research Methods 0.000 description 3
- 239000000523 sample Substances 0.000 description 3
- 238000006243 chemical reaction Methods 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 238000011161 development Methods 0.000 description 2
- 230000006872 improvement Effects 0.000 description 2
- 230000006698 induction Effects 0.000 description 2
- 238000004020 luminiscence type Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 150000002739 metals Chemical class 0.000 description 2
- 230000004044 response Effects 0.000 description 2
- 238000003466 welding Methods 0.000 description 2
- 241001479489 Peponocephala electra Species 0.000 description 1
- 230000003749 cleanliness Effects 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 239000004020 conductor Substances 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 230000006866 deterioration Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000020169 heat generation Effects 0.000 description 1
- 230000002401 inhibitory effect Effects 0.000 description 1
- 238000000691 measurement method Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000003287 optical effect Effects 0.000 description 1
- 230000000704 physical effect Effects 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 238000005070 sampling Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 238000012360 testing method Methods 0.000 description 1
- 230000005676 thermoelectric effect Effects 0.000 description 1
- 230000005514 two-phase flow Effects 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
- F23M11/00—Safety arrangements
- F23M11/04—Means for supervising combustion, e.g. windows
- F23M11/045—Means for supervising combustion, e.g. windows by observing the flame
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Photometry And Measurement Of Optical Pulse Characteristics (AREA)
Abstract
The invention discloses an ECT flame monitoring sensor attached to a combustion chamber of an aero-engine, which comprises an outer shielding layer, an outer electrode, an outer shielding electrode, an inner shielding layer, an inner electrode and an inner shielding electrode; the outer shielding layer is arranged on the outer side of the outer wall of the combustion chamber of the aeroengine and extends along the peripheral direction of the wall of the combustion chamber of the aeroengine to form a ring shape, a plurality of outer electrodes and outer shielding electrodes are arranged in the outer shielding layer, the outer electrodes are distributed along the peripheral direction of the outer shielding layer, and the outer shielding electrodes are arranged between any two adjacent outer electrodes; the inner shielding layer is arranged on the inner side of the inner wall of the central transmission shaft chamber and extends into a ring shape along the inner circumferential direction of the inner wall of the central transmission shaft chamber, a plurality of inner electrodes and inner shielding electrodes are arranged inside the inner shielding layer, the inner electrodes are distributed along the circumferential direction of the inner shielding layer, and the inner shielding electrodes are arranged between any two adjacent inner electrodes. The invention is beneficial to improving the sensitivity of the central area measurement sensitive field and improving the measurement precision.
Description
Technical Field
The invention relates to the field of aero-engine combustion chamber monitoring, in particular to an ECT flame monitoring sensor attached to an aero-engine combustion chamber.
Background
As the demand for air-borne technology increases, the thrust weight of the combustion chamber is more important than that of an aeroengine, and the distribution quality of the outlet temperature, particularly the hot spot temperature, of the combustion chamber is one of important components of the aeroengine, has a direct effect on the turbine performance and the service life of hot end components, and even can cause burning and damage of turbine-stage blades if the quality of the outlet temperature field of the combustion chamber is poor. Meanwhile, under the low emission requirement, the aero-engine is usually designed to work in a lean combustion state in consideration of comprehensive indexes such as combustion efficiency, stability and cleanliness, the air flow speed is extremely high when the aero-engine works, the residence time of fuel in a combustion chamber is in the millisecond level, so that unstable flame combustion is likely to be caused, fatigue damage is caused to the device, the reliability and the service life of the combustor are reduced, and serious and even possible engine flameout and air parking are caused, so that aero accidents are caused. Therefore, the research on the aeroengine combustion measurement technology, the real-time monitoring of the flame combustion state, has important significance for inhibiting unstable combustion, avoiding flameout and stopping and controlling and protecting the heat-sensitive components of the engine.
In the prior art, the conventional methods related to flame monitoring of the combustion chamber of the aeroengine mainly comprise the following steps:
(1) Thermocouple technology
The thermocouple is formed by welding two metals with different materials, the two metals form a loop by welding, a measuring end and a reference end are formed, the measuring end is arranged in a temperature measuring area to ensure that the measuring end is exposed in high-temperature gas flow, when the heat energy of the high-temperature gas is transferred to the temperature measuring end of the thermocouple, based on the principle of thermoelectric effect, the temperature difference between conductors with different materials can generate electromotive force, the measured electromotive force signals are transferred to a temperature conversion instrument or a control system through a thermocouple wire, and the system can calculate corresponding temperature values according to the known thermocouple characteristic curve. The thermocouple measurement mode needs to install a plurality of sensor probes to reflect the temperature and pressure changes in the flow field, when the sensors are placed in the flow field, disturbance can be caused to the flow field, measurement errors are generated, along with the improvement of the power ratio of the engine, the gas temperature at the outlet of the combustion chamber is higher and higher, the average gas temperature at the outlet of the combustion chamber is higher than 1850K, the gas temperature at the outlet of the combustion chamber is higher than 15 (more than 2000K), the measurement limit of a conventional thermocouple is exceeded, the temperature at the outlet of the combustion chamber cannot be measured by the thermocouple, and the implementation difficulty of the thermocouple technology in a visual aspect is higher.
(2) Gas analysis method
The gas analysis method is to collect high-temperature gas at the outlet of a combustion chamber, analyze the content ratio of various components in the gas, analyze the combustion reaction of the fuel by utilizing the first law of thermodynamics, directly react the components of the gas in the combustion chamber under the adiabatic condition, measure the gas components by combining parameters such as the calorific value of aviation fuel, the inlet temperature of the combustion chamber and the like, and further calculate the temperature of the gas at the outlet of the combustion chamber.
The gas analysis method is poor in instantaneity because of the need of sampling gas for testing, and the primary response time including sample gas conveying, sample gas replacement, instrument response and the like is generally about 20-30 seconds, and the gas analysis method is basically unavailable in use when the aeroplane flies at a high speed. Meanwhile, the gas analysis method is also more difficult to realize in the visual aspect of the combustion state by analyzing the combustion state and the temperature distribution through the gas components.
(3) Radiation light detection
Since the direct reaction of the fuel upon combustion is luminescence and heat generation, the hysteresis of the light is extremely small. Therefore, the detection of the state of light in the combustion chamber can quickly reflect the combustion state in the combustion chamber. By detecting the condition of the luminescence of the fuel in the combustion chamber during combustion, the ignition/flameout and flame stability of the combustion chamber of the engine can be detected. However, a relatively stable light path needs to be ensured based on radiation optical distance detection, flame combustion is performed in a flame tube for an aero-engine combustion chamber, the flame tube is used as a closed space, the light path is difficult to form, and the development of related technologies is still to be studied.
Because of the limitations of the above-described methods, ECT technology has been used. The Electric Capacitance Tomography (ECT) is a non-invasive measurement method, is commonly used in industrial two-phase flow or multiphase flow detection, and in recent years, the ECT imaging technology is rapidly developed and the application range is gradually expanded, wherein flame visual monitoring is a novel application result of the ECT technology. The ECT technology is utilized to detect the flame, so that the physical properties of the flame and the thermal radiation flow field can be ensured not to be disturbed, the flame can be monitored in real time in a visualized manner, and more comprehensive combustion information is obtained. Along with the rapid development of computer technology and the improvement of the requirements of people on detection technology, the imaging flame detection method has a very practical application prospect.
Based on ECT imaging flame detection, the capacitive tomography system (ECT) is applied to the field of flame detection by Rogerc, waterfall, university of Manchester at the earliest, and the equivalent dielectric coefficient of flame to air is measured while information such as flame position and shape is determined. Liu Dan in China, etc., accurately displays the position and shape information of the flame by using ECT and deduces and verifies a complex equivalent dielectric constant model of the flame, and further illustrates the feasibility of detecting the flame by using the Electric Capacitance Tomography (ECT) technology. Cao Zhang of Beijing aviation aerospace university and the like uses a blunt body support as an inner electrode, an annular ECT sensor is designed, and an improved Calderon method is utilized to monitor the backfire flame state of a blunt body burner, so that a new method is provided for the exploration of unstable combustion factors of flame and the visual monitoring of flame.
Currently, most of the visual monitoring researches of flame combustion state mainly comprise the combustion in an industrial boiler furnace, and the research of applying the Electric Capacitance Tomography (ECT) technology to the flame monitoring of the combustion chamber of the civil aviation engine is very rare. And for aeroengines, their combustion chambers can be broadly divided into three categories: the single-tube-shaped combustion chamber, the annular pipe-shaped combustion chamber and the full-ring combustion chamber are developed for a long time, and the full-ring combustion chamber has the advantages of compact structure, reduced weight, high space utilization rate and capability of obtaining uniform outlet temperature, so that the main position is gradually occupied, and the full-ring combustion chamber is commonly adopted by the civil aviation aircraft engine in the present stage. The annular engine combustion chamber is different from a common cavity, the chamber wall of the annular engine combustion chamber is of a metal structure, the distribution of a sensor induction field can be seriously influenced by the internal metal structure, the measurement accuracy is greatly reduced, and the annular engine combustion chamber is not suitable for measurement by using a traditional ECT circular sensor any more, so that the traditional ECT circular sensor needs to be improved.
Disclosure of Invention
The present invention aims to solve at least one of the technical problems existing in the prior art. To this end, a first object of the present invention is to provide an ECT flame-monitoring sensor that fits into the combustion chamber of an aircraft engine. The technical scheme of the invention is as follows:
an ECT flame monitoring sensor attached to a combustion chamber of an aeroengine, wherein the combustion chamber of the aeroengine is a full-ring combustion chamber and comprises an annular combustion chamber outer wall and a combustion chamber inner wall, the combustion chamber inner wall surrounds a central transmission shaft chamber for accommodating a transmission shaft, and the ECT flame monitoring sensor is arranged at an outlet of the combustion chamber of the aeroengine and is characterized by comprising an outer shielding layer, an outer electrode, an outer shielding electrode, an inner shielding layer, an inner electrode and an inner shielding electrode;
the outer shielding layer is arranged on the outer side of the outer wall of the combustion chamber of the aeroengine and extends along the peripheral direction of the wall of the combustion chamber of the aeroengine to form a ring shape, a plurality of outer electrodes and outer shielding electrodes are arranged in the outer shielding layer, the outer electrodes are distributed along the peripheral direction of the outer shielding layer, and the outer shielding electrodes are arranged between any two adjacent outer electrodes;
the inner shielding layer is arranged on the inner side of the inner wall of the central transmission shaft chamber and extends into a ring shape along the inner circumferential direction of the inner wall of the central transmission shaft chamber, a plurality of inner electrodes and inner shielding electrodes are arranged inside the inner shielding layer, the inner electrodes are distributed along the circumferential direction of the inner shielding layer, and the inner shielding electrodes are arranged between any two adjacent inner electrodes.
Further, a plurality of the external electrodes are uniformly arranged at intervals along the circumferential direction of the external shielding layer.
Further, the plurality of inner electrodes are uniformly arranged at intervals along the circumferential direction of the inner shielding layer.
Further, the total number of the external electrodes is 12, and the total number of the internal electrodes is 4.
Compared with the prior art, the invention has the following beneficial effects:
compared with the traditional ECT circular sensor, the structure of the ECT circular sensor is improved again, the ECT circular sensor is designed into the annular ECT sensor with the inner electrode and the outer electrode, the structural design of the annular ECT sensor is extremely high in geometrical conformity with the whole annular combustion chamber, meanwhile, the inner electrode and the outer electrode are also beneficial to improving the sensitivity of a central area measuring sensitive field, the measuring precision is improved, and a new idea is opened up for the flame monitoring direction of the combustion chamber of the aeroengine.
Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention.
Drawings
For a clearer description of embodiments of the invention or of solutions in the prior art, the drawings that are required to be used in the description of the embodiments or of the prior art will be briefly described, it being obvious that the drawings in the description below are only some embodiments of the invention, and that other drawings can be obtained, without the inventive effort, by a person skilled in the art from these drawings:
FIG. 1 is a schematic cross-sectional view of a prior art aircraft engine annular combustor;
FIG. 2 is a schematic illustration of the present invention disposed at an outlet of an aircraft engine combustion chamber;
fig. 3 is a usage state reference diagram of the present invention.
Reference numerals:
1. an outer wall of the combustion chamber; 2. the inner wall of the combustion chamber; 3. the outer wall of the flame tube; 4. the inner wall of the flame tube; 5. a nozzle; 6. an outer shielding layer; 7. an external electrode; 8. an outer shielding electrode; 9. an inner shielding layer; 10. an inner electrode; 11. an inner shielding electrode; 12. the inner wall of the chamber of the central transmission shaft.
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the embodiments of the present invention more apparent, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention, and it is apparent that the described embodiments are some embodiments of the present invention, but not all embodiments of the present invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
Embodiments of the present invention are described in detail below, examples of which are illustrated in the accompanying drawings, wherein like or similar reference numerals refer to like or similar elements or elements having like or similar functions throughout. In the description of the present invention, it should be understood that the terms "upper", "lower", "front", "rear", "left", "right", "inner", "outer", "vertical", "circumferential", and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, are merely for convenience in describing the present invention and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention.
In the present invention, unless explicitly specified and limited otherwise, the terms "mounted," "connected," "secured," and the like are to be construed broadly and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
As shown in fig. 1, which is a schematic cross-sectional view of a prior art annular combustion chamber for an aircraft engine, the combustion chamber is of a full annular shape, and has an annular combustion chamber outer wall 1 and a combustion chamber inner wall 2, the combustion chamber inner wall 2 meeting to define a central drive shaft chamber 12 for receiving a drive shaft. A flame tube is arranged between the outer wall and the inner wall of the combustion chamber, and a plurality of nozzles 5 which are distributed along the circumferential direction of the combustion chamber at intervals are arranged between the outer wall 3 of the flame tube and the inner wall 4 of the flame tube.
As can be seen from fig. 1, the annular engine combustion chamber is different from a common cavity, the chamber wall of the annular engine combustion chamber is of a metal structure, the distribution of the sensor induction field can be seriously affected by the internal metal structure, the measurement accuracy is greatly reduced, and the annular engine combustion chamber is not suitable for measurement by using a traditional ECT circular sensor.
Accordingly, the inventors have improved upon conventional ECT circular sensors for the particular configuration of annular combustors. As shown in fig. 2, the ECT flame-monitoring sensor includes an outer shield layer 6, an outer electrode 7, an outer shield electrode 8, an inner shield layer 9, an inner electrode 10, and an inner shield electrode 11.
The outer shielding layer 6 is arranged on the outer side of the outer wall 1 of the combustion chamber of the aeroengine and extends into a ring shape along the peripheral direction of the wall of the combustion chamber of the aeroengine, a plurality of outer electrodes 7 and outer shielding electrodes 8 are arranged inside the outer shielding layer 6, the outer electrodes 7 are uniformly distributed at intervals along the circumferential direction of the outer shielding layer 6, and the outer shielding electrodes 8 are arranged between any two adjacent outer electrodes 7.
The inner shielding layer 9 is arranged on the inner side of the inner wall 12 of the central transmission shaft chamber of the aero-engine and extends into a ring shape along the inner circumferential direction of the inner wall 12 of the central transmission shaft chamber of the aero-engine, a plurality of inner electrodes 10 and inner shielding electrodes 11 are arranged inside the inner shielding layer 9, the inner electrodes 10 are uniformly distributed at intervals along the circumferential direction of the inner shielding layer 9, and the inner shielding electrodes 11 are arranged between any two adjacent inner electrodes 10.
Of these, 12 outer electrodes 7 and 4 inner electrodes 10 are used.
As shown in fig. 3, when the ECT flame monitoring sensor in this solution is used, the ECT flame monitoring sensor needs to be disposed at the outlet of the combustion chamber of the aero-engine, so that after flame in combustion is ejected from the outlet, the flame can flow through the inner and outer electrodes 7 of the ECT flame monitoring sensor, so that the capacitance between the inner and outer electrode 7 pairs is changed, and thus the flame is monitored by the ECT flame monitoring sensor, and then relevant data processing is performed.
Compared with the traditional ECT circular sensor, the structure of the ECT circular sensor is improved again, the ECT circular sensor is designed into the annular ECT sensor with the inner electrode and the outer electrode, the structural design of the annular ECT sensor is extremely high in geometrical conformity with the whole annular combustion chamber, meanwhile, the inner electrode and the outer electrode are also beneficial to improving the sensitivity of a central area measuring sensitive field, the measuring precision is improved, and a new idea is opened up for the flame monitoring direction of the combustion chamber of the aeroengine.
The invention can realize visual monitoring of the flame state of the combustion chamber of the aeroengine and establishment of the temperature field, thereby being beneficial to solving the problems of fatigue damage, reduced reliability, service life and the like of the device of the combustion chamber of the engine caused by unstable combustion or deterioration of the outlet temperature field.
While embodiments of the present invention have been shown and described, it will be understood by those of ordinary skill in the art that: many changes, modifications, substitutions and variations may be made to the embodiments without departing from the spirit and principles of the invention, the scope of which is defined by the claims and their equivalents.
Claims (4)
1. An ECT flame monitoring sensor attached to a combustion chamber of an aeroengine, wherein the combustion chamber of the aeroengine is a full-ring combustion chamber and comprises an annular combustion chamber outer wall and a combustion chamber inner wall, the combustion chamber inner wall surrounds a central transmission shaft chamber for accommodating a transmission shaft, and the ECT flame monitoring sensor is arranged at an outlet of the combustion chamber of the aeroengine and is characterized by comprising an outer shielding layer, an outer electrode, an outer shielding electrode, an inner shielding layer, an inner electrode and an inner shielding electrode;
the outer shielding layer is arranged on the outer side of the outer wall of the combustion chamber of the aeroengine and extends along the peripheral direction of the wall of the combustion chamber of the aeroengine to form a ring shape, a plurality of outer electrodes and outer shielding electrodes are arranged in the outer shielding layer, the outer electrodes are distributed along the peripheral direction of the outer shielding layer, and the outer shielding electrodes are arranged between any two adjacent outer electrodes;
the inner shielding layer is arranged on the inner side of the inner wall of the central transmission shaft chamber and extends into a ring shape along the inner circumferential direction of the inner wall of the central transmission shaft chamber, a plurality of inner electrodes and inner shielding electrodes are arranged inside the inner shielding layer, the inner electrodes are distributed along the circumferential direction of the inner shielding layer, and the inner shielding electrodes are arranged between any two adjacent inner electrodes.
2. The ECT flame-monitoring sensor of claim 1, wherein a plurality of said outer electrodes are uniformly spaced apart along the circumference of said outer shield layer.
3. The ECT flame-monitoring sensor of claim 1, wherein a plurality of the inner electrodes are uniformly spaced apart along the circumference of the inner shield.
4. The ECT flame-monitoring sensor of claim 1, wherein there are 12 outer electrodes and 4 inner electrodes.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311701381.1A CN117646911A (en) | 2023-12-12 | 2023-12-12 | ECT flame monitoring sensor attached to aeroengine combustion chamber |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311701381.1A CN117646911A (en) | 2023-12-12 | 2023-12-12 | ECT flame monitoring sensor attached to aeroengine combustion chamber |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117646911A true CN117646911A (en) | 2024-03-05 |
Family
ID=90043211
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202311701381.1A Pending CN117646911A (en) | 2023-12-12 | 2023-12-12 | ECT flame monitoring sensor attached to aeroengine combustion chamber |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117646911A (en) |
-
2023
- 2023-12-12 CN CN202311701381.1A patent/CN117646911A/en active Pending
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CA2855035C (en) | Thermocouple | |
| CN110186583B (en) | Method for measuring temperature of ceramic matrix composite high-temperature component based on electrical impedance imaging | |
| US6062811A (en) | On-line monitor for detecting excessive temperatures of critical components of a turbine | |
| CN109341883A (en) | A kind of total temperature measurement device in aeroengine combustor buring room | |
| CN110260919B (en) | Method for simultaneously measuring temperature and strain of turbine blade tip | |
| CN215726809U (en) | Aeroengine combustion test device | |
| CN112326257A (en) | Combustion stability state monitoring and diagnosing system of gas turbine | |
| CN106872049A (en) | A kind of turbine blade surface temperature measuring equipment | |
| US9207128B2 (en) | Dynamic fiber temperature sensing package and method of assembling the same | |
| CN107328478B (en) | A method of light measurement Turbine Blade Temperature Field and emissivity are radiated based on three wave bands | |
| CN104534474B (en) | A kind of gas turbine and the method for application gas turbine detection tempering | |
| CN117646911A (en) | ECT flame monitoring sensor attached to aeroengine combustion chamber | |
| CN110146550B (en) | Method for monitoring oxidation degree of composite material high-temperature part based on electrical impedance imaging | |
| Mersinligil et al. | First unsteady pressure measurements with a fast response cooled total pressure probe in high temperature gas turbine environments | |
| CN115325567B (en) | Combustion chamber flame tube, temperature measuring system and temperature measuring method thereof | |
| CN104390716A (en) | Solid engine inflatable temperature measuring system | |
| Chana et al. | The development of a hot section eddy current sensor for turbine tip clearance measurement | |
| CN114776483B (en) | Three-dimensional transient combustion speed measuring device and combustion speed measuring method | |
| CN114088231A (en) | Temperature testing device for rotor part of low-pressure turbine in complete state of aircraft engine | |
| Rohy et al. | Radiation pyrometer for gas turbine blades | |
| Kolarević et al. | Measuring parameters of Phoenix-100 gas-generator | |
| CN106525275B (en) | A kind of multi-section acoustics tomography DC burner First air flow field analysis method | |
| Jenkins et al. | Measuring gas turbine engine component temperatures using thermographic phosphors | |
| CN216617685U (en) | Multi-surplus ignition system of aircraft engine | |
| Kolarević et al. | Crucial parameters of gas generator on tip-jet helicopter |
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 |