CN112413642A - Intelligent combustion chamber of aero-engine - Google Patents
Intelligent combustion chamber of aero-engine Download PDFInfo
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- CN112413642A CN112413642A CN202011237939.1A CN202011237939A CN112413642A CN 112413642 A CN112413642 A CN 112413642A CN 202011237939 A CN202011237939 A CN 202011237939A CN 112413642 A CN112413642 A CN 112413642A
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- optical fiber
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- combustion chamber
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
Abstract
The invention discloses an intelligent combustion chamber of an aircraft engine, which comprises a flame tube, a combustion chamber shell, a high-pressure turbine guider, a sliding arc ignition device, a laser generation beam splitting system and a combustion control system, wherein the flame tube is arranged in the combustion chamber shell, the high-pressure turbine guider and the flame tube are coaxially arranged and are positioned at an air outlet of the flame tube, and the sliding arc ignition device is arranged at the air inlet of the flame tube; the laser generation beam splitting system is arranged outside the combustion chamber shell and used for emitting measurement laser signals and comparison laser signals, the combustion control system and the laser generation beam splitting system are connected with the high-pressure turbine guider and used for transmitting the measurement laser signals, the combustion control system is further connected with the laser generation beam splitting system and used for receiving the comparison laser signals, and the combustion control system is connected with the sliding arc ignition device and used for controlling the sliding arc ignition device to work. The invention can monitor the working state of the combustion chamber of the aero-engine in real time, and adjust the working state of the combustion chamber in real time by starting the sliding arc ignition device.
Description
Technical Field
The invention relates to the field of aero-engine combustors, in particular to an aero-engine intelligent combustor capable of monitoring the working state of a flame tube of an aero-engine combustor in real time and rapidly adjusting the working state of the flame tube.
Background
The combustion chamber is one of three large core components of the aircraft engine, and the combustion quality, the combustion efficiency and the working state of the combustion chamber related to combustion are monitored and adjusted in real time, and the like, which are directly related to the safety and the economy of the aircraft engine. The working condition of the aeroengine under high-altitude flight is quite severe, the combustion chamber of the aeroengine can generate complex reaction during working, the actual combustion state is difficult to accurately display by utilizing computational fluid mechanics, and the key parameters in the combustion process of the combustion chamber can be accurately measured by an effective combustion diagnosis technology. Specifically, parameters such as the temperature of a combustion flow field on a certain interface of the combustion chamber in operation, the concentration of components of combustion products and the like need to be measured in real time, and the working state of the combustion chamber needs to be adjusted, so that the combustion chamber is always in a normal working state.
The combustion state measurement of an aircraft engine combustion chamber is divided into: the method comprises contact measurement and non-contact measurement, wherein the contact measurement can disturb the flow field distribution in the combustion chamber to influence the combustion flow field, the measurement response speed is slow, only one point of information on the section can be obtained, the information quantity is too small, the flow field of the combustion chamber is distorted, the total pressure loss of the combustion chamber is caused, and the thrust-weight ratio of the aero-engine is further influenced. Obviously, the measurement method cannot meet the requirements of acquiring the flow field information of the combustion chamber of the modern aeroengine in real time, having more data and not influencing the flow field of the combustion chamber, so that the development of a non-contact measurement technology with better transient response characteristic is required.
For the non-contact combustion chamber combustion state measurement technology, the traditional absorption spectrum measurement device comprises a prism, a grating, a Fourier transform spectrometer, a single-frequency laser and the like, the spectral resolution and the detection sensitivity are low, the frequency measurement precision and the measurement speed are not high, and the transient state measurement capability of an unstable combustion flow field needs to be improved.
Disclosure of Invention
The invention aims to overcome the defects in the prior art and provide an intelligent combustion chamber of an aero-engine, which can monitor the working state of the combustion chamber of the aero-engine in real time and adjust the working state of the combustion chamber in real time by starting a sliding arc ignition device so as to enable the combustion chamber to be always in a normal working state.
In order to achieve the purpose, the invention adopts the technical scheme that: the intelligent combustion chamber of the aircraft engine is characterized by comprising a flame tube, a combustion chamber shell, a high-pressure turbine guider, a sliding arc ignition device, a laser generation beam splitting system and a combustion control system, wherein the flame tube is arranged in the combustion chamber shell; the laser generation beam splitting system is arranged outside the combustion chamber shell and used for generating laser and splitting the generated laser into a measurement laser signal and a comparison laser signal, the high-pressure turbine guider is connected with the laser generation beam splitting system, so that the measuring laser signal enters the high-pressure turbine guider to acquire combustion flow field information at the air outlet of the flame tube, the combustion control system is connected with the high-pressure turbine guider and is used for receiving a measuring laser signal for collecting combustion flow field information at the air outlet of the flame tube, the combustion control system is connected with the laser generation beam splitting system and is used for receiving a contrast laser signal emitted by the laser generation beam splitting system, the combustion control system is connected with the sliding arc ignition device and controls the sliding arc ignition device to be opened and closed according to the received measuring laser signal and the received comparison laser signal.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the high-pressure turbine guider comprises a guider outer ring, a guider inner ring, a first guide blade, a second guide blade, a third guide blade, an incident optical fiber collimator, an intermediate optical fiber coupler, an intermediate optical fiber collimator and a receiving optical fiber coupler, wherein the first guide blade, the second guide blade and the third guide blade are sequentially fixed between the guider inner ring and the guider outer ring; the incident optical fiber collimator and the intermediate optical fiber coupler are both positioned between the first guide blade and the second guide blade, the incident optical fiber collimator is installed on the inner wall of the outer ring of the guider and close to the air inlet of the high-pressure turbine guider, the intermediate optical fiber coupler is installed on the outer wall of the inner ring of the guider and close to the air inlet of the high-pressure turbine guider, and the intermediate optical fiber coupler and the incident optical fiber collimator are positioned on the same radial section and used for receiving a measurement laser signal emitted by the incident optical fiber collimator; the receiving optical fiber coupler and the medium optical fiber collimator are both located between the second guide blade and the third guide blade, the receiving optical fiber coupler is fixedly mounted on the inner wall of the outer ring of the guider and close to the air inlet of the high-pressure turbine guider, the medium optical fiber collimator is fixedly mounted on the outer wall of the inner ring of the guider and close to the air inlet of the high-pressure turbine guider, and the receiving optical fiber coupler and the medium optical fiber collimator are located on the same radial cross section and used for receiving the measuring laser signals transmitted by the medium optical fiber collimator.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the number of the first guide blade, the second guide blade and the third guide blade is equal and is a plurality of, and the first guide blade, the second guide blade and the third guide blade are sequentially and uniformly fixed between the inner ring of the guider and the outer ring of the guider at intervals.
Foretell aeroengine intelligence combustion chamber, its characterized in that: an incident optical fiber collimator and an intermediate optical fiber coupler are uniformly arranged between each first guide blade and each second guide blade, and an intermediate optical fiber collimator and a receiving optical fiber coupler are uniformly arranged between each second guide blade and each third guide blade.
The intelligent combustion chamber of the aero-engine is characterized in that: each intermediate optical fiber coupler and the adjacent intermediate optical fiber collimator are connected through an intermediate optical fiber and used for transmitting a measuring laser signal.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the guide device outer ring and the guide device inner ring are of circular structures, and the first guide blade, the second guide blade and the third guide blade are arc-shaped blades.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the laser generation beam splitting system comprises a laser generator and a laser beam splitter, the laser beam splitter is connected with the laser generator through a number one of optical fibers and used for receiving laser emitted by the laser generator and splitting the received laser into a measurement laser signal and a contrast laser signal, and the incident optical fiber collimator is connected with the laser beam splitter through an incident optical fiber and used for receiving and emitting the measurement laser signal emitted by the laser beam splitter; the combustion control system is connected with the receiving optical fiber coupler through a receiving optical fiber and used for receiving a measuring laser signal for collecting combustion flow field information at the air outlet of the flame tube.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the combustion control system comprises a signal receiving and processing device, a main controller and a sliding arc ignition switch, wherein the signal receiving and processing device and the main controller are arranged outside a combustion chamber shell, the sliding arc ignition switch is fixedly arranged on the outer wall of the combustion chamber shell and is close to a sliding arc ignition device, the signal receiving and processing device is connected with a laser beam splitter through a second optical fiber and is used for receiving a comparison laser signal emitted by the laser beam splitter, and the signal receiving and processing device is also connected with a receiving optical fiber coupler through a receiving optical fiber and is used for receiving a measurement laser signal which collects combustion flow field information at an air outlet of a flame tube and comparing the received measurement laser signal with the comparison laser signal to perform spectral analysis; the main controller is connected with the signal receiving and processing device and used for receiving the spectral analysis signals output by the signal receiving and processing device, and the main controller is connected with the sliding arc ignition switch and used for controlling the opening and closing of the sliding arc ignition switch.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the laser generator is a tunable semiconductor laser generator.
Foretell aeroengine intelligence combustion chamber, its characterized in that: the sliding arc ignition device comprises a second swirler, a first swirler, a fuel nozzle and a venturi electrode, wherein the second swirler and the first swirler are fixedly connected with the flame tube, the first swirler is arranged in the second swirler, one end of the venturi electrode is nested outside the first swirler and is positioned between the first swirler and the second swirler, one end of the venturi electrode is connected with a power supply anode, so that the other end of the venturi electrode generates a sliding arc for igniting and supporting combustion for the flame tube, the nozzle end of the fuel nozzle is fixedly arranged in the first swirler and is used for injecting oil into the flame tube, the other end of the fuel nozzle is connected with a power supply cathode, and the other end of the venturi electrode extends into the flame tube.
Compared with the prior art, the invention has the following advantages:
1. structurally, the incident optical fiber collimator and the receiving optical fiber coupler are arranged on the outer ring of the guider at the air inlet of the high-pressure turbine guider at intervals in a staggered mode along the circumferential direction of the outer ring of the high-pressure turbine guider and are used for measuring the emission and the receiving of laser signals respectively, the laser signals are arranged regularly to form a blocking surface of airflow, the blocking surface covers the cross section of an air outlet of a flame tube of a combustion chamber in a large area, and the measured data are sufficient. The optical fiber does not enter a combustion flow field, so that total pressure loss caused by smooth combustion distortion can be avoided.
2. In terms of working principle, the signal receiving and processing device compares the received comparison laser signal with the received measurement laser signal and performs spectral analysis, information such as combustion flow field temperature, combustion product component concentration and the like is fed back to the main controller, the main controller judges the current working state of the combustion chamber according to the received spectral analysis and adjusts the working state of the combustion chamber by starting the sliding arc ignition device, and the working state of the combustion chamber of the aero-engine can be monitored and adjusted in real time.
3. In the ignition mode, the sliding arc ignition mode is adopted, the main controller controls the sliding arc to ignite, active particles generated during the discharge of the sliding arc participate in the combustion reaction to realize the combustion supporting effect, and the combustion quality is improved. The sliding arc can also crack kerosene particles, break the carbon chain of high-carbon molecules to form low-carbon micromolecules, and is more favorable for ignition. Since the area of the sliding arc discharge is located between the cathode fuel nozzle and the anode venturi, the fuel spray is immediately ignited through the discharge area, thereby igniting the entire combustion chamber.
4. The sliding arc ignition device can directly replace the existing combustion chamber ignition device in structure, can realize ignition and combustion supporting of the combustion chamber without complete replacement, is started when the working condition of the combustion chamber is poor, realizes the duty function, and ensures that the combustion chamber does not flameout.
The invention is described in further detail below with reference to the figures and examples.
Drawings
FIG. 1 is a schematic view of the combustor with a partial combustor casing removed.
FIG. 2 is a schematic view of a portion of the structure of the present invention including the point A in FIG. 1.
FIG. 3 is a schematic view of the high pressure turbine nozzle at A in FIG. 1.
Fig. 4 is a schematic view of the sliding arc ignition device of the present invention.
Fig. 5 is a schematic block diagram of the circuit of the present invention.
Description of reference numerals:
1-sliding arc ignition device; 1-a fuel nozzle; 1-2-a second cyclone;
1-3-a venturi electrode; 1-4-a first cyclone; 2, a flame tube;
3-laser generator; 4-laser beam splitter; 5-a combustion chamber housing;
6-incident optical fiber; 7-receiving optical fiber; 8-high pressure turbine guide;
8-1-a director outer ring; 8-2-guide vane number one; 8-3-second guide vane;
8-4-incident fiber collimator; 8-5-a receiving fiber coupler; 8-6-third guide vane;
8-7-intermediate optical fiber collimator; 8-intermediate fiber coupler; 8-9-guide inner ring;
9-an intermediate optical fiber; 10-optical fiber number one; 11-fiber number two;
12-a signal receiving and processing unit; 13 — a master controller; 14-sliding arc ignition switch.
Detailed Description
As shown in fig. 1 and fig. 2, the invention includes a flame tube 2, a combustion chamber housing 5, a high-pressure turbine guider 8, a sliding arc ignition device 1, a laser generation beam splitting system and a combustion control system, wherein the flame tube 2 is arranged in the combustion chamber housing 5, the high-pressure turbine guider 8 is arranged at an air outlet of the flame tube 2, the high-pressure turbine guider 8 and the flame tube 2 are coaxially arranged, and the sliding arc ignition device 1 is arranged at the air inlet of the flame tube 2 and realizes ignition combustion supporting for the flame tube 2; the laser generation beam splitting system is arranged outside the combustion chamber shell 5 and used for generating laser and splitting the generated laser into a measuring laser signal and a comparison laser signal, the high-pressure turbine guider 8 is connected with the laser generation beam splitting system, so that the measuring laser signal enters the high-pressure turbine guider 8 to collect combustion flow field information at the air outlet of the flame tube 2, the combustion control system is connected with the high-pressure turbine guider 8 and is used for receiving a measuring laser signal for collecting combustion flow field information at the air outlet of the flame tube 2, the combustion control system is connected with the laser generation beam splitting system and is used for receiving a contrast laser signal emitted by the laser generation beam splitting system, the combustion control system is connected with the sliding arc ignition device 1 and controls the opening and closing of the sliding arc ignition device 1 by analyzing the received measuring laser signal and the received comparison laser signal.
As shown in fig. 3, the high-pressure turbine guider 8 comprises a guider outer ring 8-1, a guider inner ring 8-9, a first guide blade 8-2, a second guide blade 8-3, a third guide blade 8-6, an incident optical fiber collimator 8-4, an intermediate optical fiber coupler 8-8, an intermediate optical fiber collimator 8-7 and a receiving optical fiber coupler 8-5, wherein the first guide blade 8-2, the second guide blade 8-3 and the third guide blade 8-6 are sequentially fixed between the guider inner ring 8-9 and the guider outer ring 8-1; the incident optical fiber collimator 8-4 and the intermediate optical fiber coupler 8-8 are located between the first guide blade 8-2 and the second guide blade 8-3, the incident optical fiber collimator 8-4 is installed on the inner wall of the guide outer ring 8-1 and close to an air inlet of the high-pressure turbine guide 8, the intermediate optical fiber coupler 8-8 is installed on the outer wall of the guide inner ring 8-9 and close to the air inlet of the high-pressure turbine guide 8, and the intermediate optical fiber coupler 8-8 and the incident optical fiber coupler 8-4 are located on the same radial section and used for receiving a measurement laser signal emitted by the incident optical fiber collimator 8-4; the receiving optical fiber coupler 8-5 and the intermediate optical fiber collimator 8-7 are located between the second guide blade 8-3 and the third guide blade 8-6, the receiving optical fiber coupler 8-5 is fixedly installed on the inner wall of the outer ring 8-1 of the guider and close to an air inlet of the high-pressure turbine guider 8, the intermediate optical fiber collimator 8-7 is fixedly installed on the outer wall of the inner ring 8-9 of the guider and close to the air inlet of the high-pressure turbine guider 8, and the receiving optical fiber coupler 8-5 and the intermediate optical fiber collimator 8-7 are located on the same radial cross section and used for receiving measuring laser signals transmitted by the intermediate optical fiber collimator 8-7.
As shown in fig. 3, the number of the first guide blade 8-2, the second guide blade 8-3 and the third guide blade 8-6 is equal and multiple, and the multiple first guide blades 8-2, the second guide blades 8-3 and the third guide blades 8-6 are sequentially and uniformly fixed between the guide inner ring 8-9 and the guide outer ring 8-1 at intervals.
An incident optical fiber collimator 8-4 and an intermediate optical fiber coupler 8-8 are uniformly arranged between each first guide blade 8-2 and each second guide blade 8-3, and an intermediate optical fiber collimator 8-7 and a receiving optical fiber coupler 8-5 are uniformly arranged between each second guide blade 8-3 and each third guide blade 8-6.
Each of the intermediate fiber couplers 8-8 and an adjacent one of the intermediate fiber collimators 8-7 are connected through an intermediate fiber 9 for transmitting a measuring laser signal.
The guide device outer ring 8-1 and the guide device inner ring 8-9 are both of circular structures, and the first guide blade 8-2, the second guide blade 8-3 and the third guide blade 8-6 are all arc-shaped blades.
As shown in fig. 2, the laser generation and beam splitting system includes a laser generator 3 and a laser beam splitter 4, the laser beam splitter 4 is connected with the laser generator 3 through a first optical fiber 10 for receiving laser emitted by the laser generator 3 and splitting the received laser into a measurement laser signal and a comparison laser signal, and an incident optical fiber collimator 8-4 is connected with the laser beam splitter 4 through an incident optical fiber 6 for receiving and emitting the measurement laser signal emitted by the laser beam splitter 4; the combustion control system is connected with the receiving optical fiber coupler 8-5 through the receiving optical fiber 7 and is used for receiving the measuring laser signal for collecting the combustion flow field information at the air outlet of the flame tube 2.
As shown in fig. 3, a plurality of the incident optical fiber collimators 8-4 and the receiving optical fiber couplers 8-5 are distributed along the circumference of the guide outer ring 8-1, and the incident optical fiber collimators 8-4 and the receiving optical fiber couplers 8-5 are arranged at intervals in a staggered manner; a plurality of intermediate optical fiber couplers 8-8 and intermediate optical fiber collimators 8-7 are circumferentially distributed along the inner ring 8-9 of the guider, and the intermediate optical fiber couplers 8-8 and the intermediate optical fiber collimators 8-7 are arranged at intervals in a staggered manner; the number of the incident optical fiber collimator 8-4, the receiving optical fiber coupler 8-5, the intermediate optical fiber collimator 8-7 and the intermediate optical fiber coupler 8-8 are equal.
In the embodiment, the incident optical fiber collimator 8-4 and the receiving optical fiber coupler 8-5 are regularly arranged on the guider outer ring 8-1 at the air inlet of the high-pressure turbine guider 8 along the same radial section of the combustion chamber and are respectively used for measuring the emission and the reception of laser signals, the measuring laser is regularly arranged to form a blocking surface of airflow, the blocking surface covers the section of the combustion chamber outlet in a large area, and the measuring data is sufficient. The optical fiber does not enter the combustion flow field, so that the total pressure loss caused by the flow field distortion in the combustion chamber can be avoided.
As shown in fig. 5, the combustion control system includes a signal receiving and processing unit 12, a main controller 13 and a sliding arc ignition switch 14, the signal receiving and processing unit 12 and the main controller 13 are both disposed outside the combustion chamber housing 5, the sliding arc ignition switch 14 is fixedly mounted on the outer wall of the combustion chamber housing 5 and near the sliding arc ignition device 1, the signal receiving and processing unit 12 is connected with the laser beam splitter 4 through a second optical fiber 11 for receiving a comparison laser signal emitted by the laser beam splitter 4, the signal receiving and processing unit 12 is further connected with a receiving optical fiber coupler 8-5 through a receiving optical fiber 7 for receiving a measurement laser signal that collects combustion flow field information at the air outlet of the flame tube 2 and comparing the received measurement laser signal with the comparison laser signal for making a spectrum analysis; the main controller 13 is connected with the signal receiving and processing unit 12 and used for receiving the spectral analysis signal output by the signal receiving and processing unit 12, and the main controller 13 is connected with the sliding arc ignition switch 14 and used for controlling the on and off of the sliding arc ignition switch 14.
The signal receiving and processing unit 12 compares the peak value, narrow line width and wavelength of the received measurement laser signal with the peak value, narrow line width and wavelength corresponding to the comparison laser signal, and then obtains the temperature of the combustion flow field in the flame tube 2 and the concentration information of the combustion product components through calculation, wherein the concentration information of the combustion product components includes the concentration information of water molecules, the concentration information of carbon oxides and the concentration information of nitrogen oxides of the combustion product. The main controller 13 is connected with the signal receiving and processing unit 12 and is used for receiving combustion flow field information and judging the current working state of the combustion chamber; the main controller 13 is connected with the sliding arc ignition switch 14 and is used for controlling the sliding arc ignition switch 14 to start the venturi tube electrode 1-3 in the sliding arc ignition device 1 to discharge and controlling the fuel injection quantity of the fuel nozzle 1-1, so that the working state of the combustion chamber is adjusted.
In this embodiment, the laser generator 3 is a tunable semiconductor laser generator.
As shown in fig. 4, the sliding arc ignition device 1 comprises a second swirler 1-2, a first swirler 1-4, a fuel nozzle 1-1 and a venturi electrode 1-3, wherein the second swirler 1-2 and the first swirler 1-4 are both fixedly connected with the flame tube 2, the first swirler 1-4 is arranged in the second swirler 1-2, one end of the venturi electrode 1-3 is nested outside the first swirler 1-4 and is positioned between the first swirler 1-4 and the second swirler 1-2, one end of the venturi electrode 1-3 is connected with a power supply anode so that the other end of the venturi electrode 1-3 generates a sliding arc for igniting and supporting combustion for the flame tube, the nozzle end of the fuel nozzle 1-1 is fixedly arranged in the first swirler 1-4 for injecting oil into the flame tube 2, the other end of the fuel nozzle 1-1 is connected with a power supply cathode, and the other end of the venturi electrode 1-3 extends into the flame tube 2.
The signal receiving and processing unit 12 is based on the TDLAS temperature measurement technique, and can tune the semiconductor laser absorption spectrum temperature measurement technique, and its theory of operation is: the characteristic that the narrow line width and the wavelength of the tunable semiconductor laser change along with the injection current is utilized to realize the measurement of the absorption light of a single molecule or a plurality of molecules which are very close and difficult to distinguish, the signal receiving and processing device 12 compares the received comparison laser signal with the measurement laser signal and makes spectral analysis, and the temperature of the measured gas and the concentration parameters of the combustion product components are accurately obtained.
The intelligent combustion chamber of the aero-engine adopts a novel ignition mode based on sliding arc ignition, combustion supporting of the combustion chamber can be achieved, and combustion quality and combustion efficiency are improved. The sliding arc ignition device can directly replace the existing ignition structure, ignition and combustion supporting of a combustion chamber can be realized without complete replacement, a sliding arc discharge area is positioned between a fuel nozzle and a venturi tube electrode, fuel spray can be immediately ignited through the discharge area, and then the whole combustion chamber is ignited. Active particles generated in the sliding arc discharge process participate in the combustion reaction to realize the combustion supporting effect and improve the combustion quality. Meanwhile, the sliding arc can also crack kerosene particles, so that carbon chains of high-carbon molecules are broken to form low-carbon small molecules, and ignition is facilitated.
The working principle of the invention is as follows: the laser generated by the laser generator 3 is transmitted to the laser beam splitter 4 through the first optical fiber 10, the laser generated by the laser generator 3 is divided into at least two paths of laser by the laser beam splitter 4, wherein one path of laser is used as a contrast laser signal and is transmitted to the signal receiving and processing unit 12 through the second optical fiber 11, and the rest of laser is used as a measurement laser signal and is transmitted to the incident optical fiber collimator 8-4 fixedly arranged on the guider outer ring 8-1 through the incident optical fiber 6.
The measurement laser signal is directly emitted into a combustion flow field by the incident optical fiber collimator 8-4 without being transmitted by an optical fiber to acquire information such as temperature and concentration of combustion products at the gas outlet of the flame tube 2, and then the measurement laser signal is received by the intermediary optical fiber coupler 8-8, the intermediary optical fiber coupler 8-8 is connected with the intermediary optical fiber collimator 8-7 through an intermediary optical fiber 9 to transmit the measurement laser signal, the measurement laser signal is input into the intermediary optical fiber collimator 8-7 through the intermediary optical fiber 9 and then emitted by the emitting end of the intermediary optical fiber collimator 8-7, the measurement laser signal is emitted into the combustion flow field again, and then the measurement laser signal is received by the receiving optical fiber coupler 8-5. The receiving optical fiber coupler 8-5 transmits the received measuring laser signal to the signal receiving and processing device 12 through the receiving optical fiber 7, and the signal receiving and processing device 12 compares the peak value, narrow line width and wavelength of the received measuring laser signal with the peak value, narrow line width and wavelength corresponding to the comparison laser signal and makes a spectrum analysis to accurately obtain the information of the temperature of the measured gas, the component concentration of the combustion product and the like.
The main controller 13 is connected with the signal receiving and processing unit 12 and is used for receiving the spectral analysis result input by the signal receiving and processing unit 12 and judging the current working state of the combustion chamber, if the temperature in the flame tube is measured to be lower than the set value of 100 ℃, the main controller 13 controls the sliding arc ignition switch 14 to start the Venturi tube electrode 1-3 in the sliding arc ignition device 1 to discharge to generate a sliding arc and increase the fuel injection quantity of the fuel nozzle 1-1 until the temperature reaches the normal temperature value. When the aero-engine is in a bad working condition and the combustion is insufficient, the concentration of the combustion product components is reduced, the main controller 13 commands the sliding arc ignition switch 14 to start the Venturi electrodes 1-3 in the sliding arc ignition device 1 to discharge, the combustion is strengthened, and the combustion quality is improved.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and all simple modifications, changes and equivalent structural changes made to the above embodiment according to the technical spirit of the present invention still fall within the protection scope of the technical solution of the present invention.
Claims (10)
1. The utility model provides an aeroengine intelligence combustion chamber which characterized in that: the laser combustion chamber comprises a flame tube (2), a combustion chamber shell (5), a high-pressure turbine guider (8), a sliding arc ignition device (1), a laser generation beam splitting system and a combustion control system, wherein the flame tube (2) is arranged in the combustion chamber shell (5), the high-pressure turbine guider (8) is arranged at an air outlet of the flame tube (2), the high-pressure turbine guider (8) and the flame tube (2) are coaxially arranged, and the sliding arc ignition device (1) is arranged at the air inlet of the flame tube (2) and is used for realizing ignition and combustion supporting on the flame tube (2); the laser generation beam splitting system is arranged outside the combustion chamber shell (5) and is used for generating laser and splitting the generated laser into a measurement laser signal and a comparison laser signal, the high-pressure turbine guider (8) is connected with the laser generation beam splitting system, so that the measuring laser signal enters the high-pressure turbine guider (8) to collect combustion flow field information at the air outlet of the flame tube (2), the combustion control system is connected with the high-pressure turbine guider (8) and is used for receiving a measuring laser signal for collecting combustion flow field information at the air outlet of the flame tube (2), the combustion control system is connected with the laser generation beam splitting system and is used for receiving a contrast laser signal emitted by the laser generation beam splitting system, the combustion control system is connected with the sliding arc ignition device (1) and controls the opening and closing of the sliding arc ignition device (1) by analyzing the received measuring laser signal and the comparison laser signal.
2. The aircraft engine intelligent combustion chamber of claim 1, wherein: the high-pressure turbine guider (8) comprises a guider outer ring (8-1), a guider inner ring (8-9), a first guide blade (8-2), a second guide blade (8-3), a third guide blade (8-6), an incident optical fiber collimator (8-4), an intermediate optical fiber coupler (8-8), an intermediate optical fiber collimator (8-7) and a receiving optical fiber coupler (8-5), wherein the first guide blade (8-2), the second guide blade (8-3) and the third guide blade (8-6) are sequentially fixed between the guider inner ring (8-9) and the guider outer ring (8-1); the incident optical fiber collimator (8-4) and the intermediate optical fiber coupler (8-8) are located between the first guide blade (8-2) and the second guide blade (8-3), the incident optical fiber collimator (8-4) is installed on the inner wall of the guide device outer ring (8-1) and close to an air inlet of the high-pressure turbine guide device (8), the intermediate optical fiber coupler (8-8) is installed on the outer wall of the guide device inner ring (8-9) and close to the air inlet of the high-pressure turbine guide device (8), and the intermediate optical fiber coupler (8-8) and the incident optical fiber collimator (8-4) are located on the same radial section and used for receiving a measurement laser signal emitted by the incident optical fiber collimator (8-4); the receiving optical fiber coupler (8-5) and the intermediate optical fiber collimator (8-7) are located between the second guide blade (8-3) and the third guide blade (8-6), the receiving optical fiber coupler (8-5) is fixedly installed on the inner wall of the outer guide ring (8-1) and close to the air inlet of the high-pressure turbine guide (8), the intermediate optical fiber collimator (8-7) is fixedly installed on the outer wall of the inner guide ring (8-9) and close to the air inlet of the high-pressure turbine guide (8), and the receiving optical fiber coupler (8-5) and the intermediate optical fiber collimator (8-7) are located on the same radial section and used for receiving a measuring laser signal emitted by the intermediate optical fiber collimator (8-7).
3. The aircraft engine intelligent combustion chamber of claim 2, wherein: the number of the first guide blade (8-2), the second guide blade (8-3) and the third guide blade (8-6) is equal and is multiple, and the first guide blade (8-2), the second guide blade (8-3) and the third guide blade (8-6) are sequentially and uniformly fixed between the inner ring (8-9) of the guider and the outer ring (8-1) of the guider at intervals.
4. An aircraft engine intelligent combustion chamber according to claim 3, characterized in that: an incident optical fiber collimator (8-4) and an intermediate optical fiber coupler (8-8) are uniformly arranged between each first guide blade (8-2) and each second guide blade (8-3), and an intermediate optical fiber collimator (8-7) and a receiving optical fiber coupler (8-5) are uniformly arranged between each second guide blade (8-3) and each third guide blade (8-6).
5. An aircraft engine intelligent combustion chamber according to any one of claims 1 to 4, characterised in that: each intermediate fiber coupler (8-8) and an adjacent intermediate fiber collimator (8-7) are connected through an intermediate fiber (9) for transmitting a measuring laser signal.
6. An aircraft engine intelligent combustion chamber according to any one of claims 1 to 4, characterised in that: the guide device outer ring (8-1) and the guide device inner ring (8-9) are both of a circular structure, and the first guide blade (8-2), the second guide blade (8-3) and the third guide blade (8-6) are all arc-shaped blades.
7. The aircraft engine intelligent combustion chamber of claim 2, wherein: the laser generation beam splitting system comprises a laser generator (3) and a laser beam splitter (4), the laser beam splitter (4) is connected with the laser generator (3) through a first optical fiber (10) and is used for receiving laser emitted by the laser generator (3) and dividing the received laser into a measurement laser signal and a comparison laser signal, and an incidence optical fiber collimator (8-4) is connected with the laser beam splitter (4) through an incidence optical fiber (6) and is used for receiving and emitting the measurement laser signal emitted by the laser beam splitter (4); the combustion control system is connected with the receiving optical fiber coupler (8-5) through a receiving optical fiber (7) and is used for receiving a measuring laser signal for collecting combustion flow field information at an air outlet of the flame tube (2).
8. The aircraft engine intelligent combustion chamber of claim 7, wherein: the combustion control system comprises a signal receiving and processing device (12), a main controller (13) and a sliding arc ignition switch (14), the signal receiving and processing unit (12) and the main controller (13) are both arranged outside the combustion chamber shell (5), the sliding arc ignition switch (14) is fixedly arranged on the outer wall of the combustion chamber shell (5) and is close to the sliding arc ignition device (1), the signal receiving and processing unit (12) is connected with the laser beam splitter (4) through a second optical fiber (11) and is used for receiving a contrast laser signal emitted by the laser beam splitter (4), the signal receiving and processing device (12) is also connected with the receiving optical fiber coupler (8-5) through a receiving optical fiber (7) and is used for receiving a measuring laser signal for collecting combustion flow field information at the air outlet of the flame tube (2) and comparing the received measuring laser signal with a comparison laser signal to perform spectral analysis; the main controller (13) is connected with the signal receiving and processing unit (12) and used for receiving the spectral analysis signals output by the signal receiving and processing unit (12), and the main controller (13) is connected with the sliding arc ignition switch (14) and used for controlling the opening and closing of the sliding arc ignition switch (14).
9. The aircraft engine intelligent combustion chamber of claim 7, wherein: the laser generator (3) is a tunable semiconductor laser generator.
10. The aircraft engine intelligent combustion chamber of claim 6, wherein: the sliding arc ignition device (1) comprises a second swirler (1-2), a first swirler (1-4), a fuel nozzle (1-1) and a venturi electrode (1-3), wherein the second swirler (1-2) and the first swirler (1-4) are fixedly connected with a flame tube (2), the first swirler (1-4) is arranged in the second swirler (1-2), one end of the venturi electrode (1-3) is nested outside the first swirler (1-4) and is positioned between the first swirler (1-4) and the second swirler (1-2), one end of the venturi electrode (1-3) is connected with a power supply anode, so that the other end of the venturi electrode (1-3) generates a sliding arc for igniting and supporting combustion of the flame tube, the nozzle end of the fuel nozzle (1-1) is nested in the first swirler (1-4) and used for injecting oil into the flame tube (2), the other end of the fuel nozzle (1-1) is connected with a power supply cathode, and the other end of the venturi electrode (1-3) extends into the flame tube (2).
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113898974A (en) * | 2021-10-19 | 2022-01-07 | 中国人民解放军空军工程大学 | Aeroengine combustion chamber sliding arc plasma on-duty flame head |
| CN115325567A (en) * | 2022-07-05 | 2022-11-11 | 中国航发湖南动力机械研究所 | Combustion chamber flame tube and temperature measuring system and method thereof |
| CN115875136A (en) * | 2023-01-16 | 2023-03-31 | 北京航科星云科技有限公司 | Sliding arc detection and adjustment method and system |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6294389B1 (en) * | 1999-11-01 | 2001-09-25 | General Motors Corporation | Furnace for an engine exhaust measurement system |
| CN107366928A (en) * | 2017-06-20 | 2017-11-21 | 中国人民解放军空军工程大学 | A kind of Plasma Assisted Combustion method of engine chamber |
| CN109668169A (en) * | 2018-12-20 | 2019-04-23 | 中国人民解放军空军工程大学 | A kind of aeroengine combustor buring room plasmaassisted is atomized ignition burner |
| CN110657993A (en) * | 2019-10-17 | 2020-01-07 | 北京航空航天大学 | Method for monitoring combustion field of aero-engine based on all-fiber optical frequency comb system |
| CN111206964A (en) * | 2018-11-22 | 2020-05-29 | 中发天信(北京)航空发动机科技股份有限公司 | Integrally cast aeroengine turbine guider and preparation method thereof |
| CN111795829A (en) * | 2020-07-07 | 2020-10-20 | 西安热工研究院有限公司 | TDLAS technology-based gas turbine combustion state parameter monitoring system |
-
2020
- 2020-11-09 CN CN202011237939.1A patent/CN112413642B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6294389B1 (en) * | 1999-11-01 | 2001-09-25 | General Motors Corporation | Furnace for an engine exhaust measurement system |
| CN107366928A (en) * | 2017-06-20 | 2017-11-21 | 中国人民解放军空军工程大学 | A kind of Plasma Assisted Combustion method of engine chamber |
| CN111206964A (en) * | 2018-11-22 | 2020-05-29 | 中发天信(北京)航空发动机科技股份有限公司 | Integrally cast aeroengine turbine guider and preparation method thereof |
| CN109668169A (en) * | 2018-12-20 | 2019-04-23 | 中国人民解放军空军工程大学 | A kind of aeroengine combustor buring room plasmaassisted is atomized ignition burner |
| CN110657993A (en) * | 2019-10-17 | 2020-01-07 | 北京航空航天大学 | Method for monitoring combustion field of aero-engine based on all-fiber optical frequency comb system |
| CN111795829A (en) * | 2020-07-07 | 2020-10-20 | 西安热工研究院有限公司 | TDLAS technology-based gas turbine combustion state parameter monitoring system |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113898974A (en) * | 2021-10-19 | 2022-01-07 | 中国人民解放军空军工程大学 | Aeroengine combustion chamber sliding arc plasma on-duty flame head |
| CN113898974B (en) * | 2021-10-19 | 2022-10-04 | 中国人民解放军空军工程大学 | Aero-engine combustion chamber sliding arc plasma on-duty flame head |
| CN115325567A (en) * | 2022-07-05 | 2022-11-11 | 中国航发湖南动力机械研究所 | Combustion chamber flame tube and temperature measuring system and method thereof |
| CN115875136A (en) * | 2023-01-16 | 2023-03-31 | 北京航科星云科技有限公司 | Sliding arc detection and adjustment method and system |
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