Method for diagnosing combustion instability excitation source through acoustic matrix of rocket engine
By uniformly arranging acoustic pressure sensors on rocket engines and applying independent component analysis and FFT technology, the problem of identifying combustion instability excitation sources has been solved, achieving high-precision combustion diagnosis and improved safety.
Patent Information
- Application Number
- CN202510431855.8
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-04-08
- Publication Date
- 2025-08-01
AI Technical Summary
Existing technologies struggle to accurately identify the excitation sources of combustion instability within rocket engine combustion chambers, leading to difficulties in combustion diagnosis and impacting engine performance and structural safety.
Multiple acoustic pressure sensors are uniformly arranged along the axial and circumferential directions of the rocket engine. The acoustic pressure signals are processed by independent component analysis and combined with FFT technology to determine the characteristic frequency and the dominant frequency, thereby accurately locating the excitation source.
It enables accurate positioning of combustion instability excitation sources, improves the accuracy and safety of combustion diagnosis, is applicable to various engine structures, and is unaffected by high temperature and high pressure environments.
Smart Images
Figure CN120402253A_ABST
Abstract
Description
Technical Field
[0001] This application relates to the field of engine fault diagnosis. Specifically, it relates to a method for diagnosing the excitation source of combustion instability in a rocket engine using an acoustic matrix. Background Technique
[0002] For a long time, the problem of combustion instability that occurs during the development of liquid rocket engines has been considered one of the most difficult problems. When the thermal component - the thrust chamber experiences unstable combustion, the engine performance will drop significantly. Among the set detection signals, the more intuitive characteristic of the time-series signal is that the amplitude of the combustion chamber pressure is 10% - 1000% of the steady-state chamber pressure. When performing a fast Fourier transform on the dynamic pressure signal, obvious frequency spikes will appear in its frequency spectrum distribution, and the amplitude of the frequency spikes is much higher than that of other frequencies. When there are large pressure oscillations in the combustion chamber, it is extremely easy to cause serious consequences such as structural damage. Therefore, in the engine development stage, in-depth experimental research on unstable combustion needs to be carried out. When the entire engine is running, for the thermal component - the thrust chamber, there are multiple excitation sources for combustion instability, that is, triggering causes. Typical excitation source triggering mechanisms are: (1) When the upstream propellant flows through the pump, its flow pulsation frequency is easily coupled with the combustion eigenfrequency in the downstream combustion chamber, which will form a positive excitation, resulting in the generation of combustion instability; (2) When the propellant in the engine flows through complex pipelines, the frequency formed in the pipeline has a certain positive amplification effect on the pulsation of the propellant. When it is transmitted downstream, it will form an initial excitation source, and in the combustion chamber, this perturbation may be further amplified, resulting in the generation of combustion instability; (3) Different components in the engine will form independent acoustic cavities in each chamber. After passing through chambers with different volumes and structures, the pressure pulsation frequency will also change accordingly. As the propellant flows, various pressure pulsation frequencies will be transmitted downstream, forming perturbation factors for the combustion chamber. Different excitation sources have their own obvious acoustic characteristics, but in a complex thermal environment, various sounds will be mixed together, making it difficult to distinguish. Therefore, in the development stage, determining which excitation source in the combustion chamber triggers combustion instability, that is, accurate combustion diagnosis and positioning, is the premise for the design improvement direction.
[0003] Currently, the signal monitoring methods in hot fire tests are relatively single. Commonly used are vibration pressure sensors and dynamic pressure sensors. Although these two signals can monitor the time-series signals at various parts of the engine, due to the differences in the measurement characteristics and installation methods of the two sensors, it is very difficult for the two signals to form a comprehensive coupling analysis. Summary of the Invention
[0004] To overcome at least one deficiency in the prior art, this application provides a method for diagnosing the excitation source of combustion instability in a rocket engine using an acoustic matrix.
[0005] In a first aspect, a method for diagnosing the excitation source of combustion instability in a rocket engine using an acoustic matrix is provided, including:
[0006] Acquiring sound pressure signals monitored by a plurality of sound pressure sensors uniformly arranged along the axial and circumferential directions of the rocket engine;
[0007] All sound pressure signals are processed using independent component analysis to obtain multiple sound pressure signals with different characteristic frequencies;
[0008] Performing FFT on a plurality of sound pressure signals with different characteristic frequencies to determine the characteristic frequency of each sound pressure signal;
[0009] Obtaining the main frequency monitored in the engine combustion chamber and determining the sound pressure signal corresponding to the characteristic frequency identical to the main frequency as the target sound pressure signal;
[0010] The sound pressure sensor arrangement position corresponding to the target sound pressure signal is the position where the excitation source is finally determined.
[0011] In one embodiment, a plurality of acoustic pressure sensors are evenly arranged along the axial direction of the rocket engine, including:
[0012] They are evenly arranged from the top of the rocket engine to the nozzle outlet.
[0013] In a second aspect, a device for diagnosing combustion instability excitation sources using an acoustic matrix of a rocket engine is provided, comprising:
[0014] A sound pressure signal acquisition module is used to acquire sound pressure signals monitored by multiple sound pressure sensors evenly arranged along the axial and circumferential directions of the rocket engine;
[0015] A signal processing module is used to process all sound pressure signals using an independent component analysis method to obtain multiple sound pressure signals with different characteristic frequencies;
[0016] a characteristic frequency determination module, configured to perform FFT on a plurality of sound pressure signals having different characteristic frequencies to determine the characteristic frequency of each sound pressure signal;
[0017] a target sound pressure signal determination module, configured to obtain the main frequency monitored in the engine combustion chamber and determine a sound pressure signal corresponding to a characteristic frequency identical to the main frequency as the target sound pressure signal;
[0018] The excitation source position determination module is used to determine that the sound pressure sensor arrangement position corresponding to the target sound pressure signal is the position where the excitation source is finally determined.
[0019] In one embodiment, a plurality of acoustic pressure sensors are evenly arranged along the axial direction of the rocket engine, including:
[0020] They are evenly arranged from the top of the rocket engine to the nozzle outlet.
[0021] In a third aspect, a system for diagnosing the combustion instability excitation source of a rocket engine using an acoustic matrix is provided, including: a rocket engine, a data transmission line, a plurality of sound pressure sensors, a data receiver, and a data processor;
[0022] The plurality of sound pressure sensors are uniformly arranged along the axial and circumferential directions of the rocket engine for detecting the sound pressure signals at different positions;
[0023] The sound pressure signals at different positions are sent to the data receiver through the data transmission line;
[0024] The data receiver sends the sound pressure signals at different positions to the data processor; the data processor is used to implement the method for diagnosing the combustion instability excitation source of the rocket engine using the acoustic matrix as described above.
[0025] Compared with the prior art, the present application has the following beneficial effects: The method for diagnosing the combustion instability excitation source of the rocket engine using the acoustic matrix of the present application realizes the acquisition of the sound pressure at different positions of the engine. The collected signal data is analyzed by the ICA method, and the parts and frequencies where the combustion instability excitation source of the engine is concentrated can be obtained. The method of the present application has the advantages of cross-model, non-contact, and reusable. BRIEF DESCRIPTION OF THE DRAWINGS
[0026] The present application can be better understood by referring to the description given below in conjunction with the accompanying drawings. The drawings, together with the following detailed description, are included in this specification and form a part of this specification. In the drawings:
[0027] Figure 1 Shows a flowchart of the method for diagnosing the combustion instability excitation source of the rocket engine using the acoustic matrix;
[0028] Figure 2 Shows a schematic diagram of the arrangement of a plurality of sound pressure sensors along the axial direction of the rocket engine;
[0029] Figure 3 Shows a schematic diagram of the arrangement of a plurality of sound pressure sensors along the circumferential direction of the rocket engine.
[0030] Reference Signs:
[0031] 1 - data transmission line; 2 - sound pressure sensor; 3 - rocket engine; 4 - data receiver; 5 - data processor. DETAILED DESCRIPTION OF THE EMBODIMENTS
[0032] Exemplary embodiments of the present application will be described below in conjunction with the accompanying drawings. For clarity and conciseness, not all features of the actual embodiments are described in the specification. However, it should be understood that many embodiment-specific decisions may be made during the development of any such actual embodiment to achieve the specific goals of the developer, and these decisions may vary with different embodiments.
[0033] Here, it should also be noted that in order to avoid obscuring the present application with unnecessary details, only the device structures closely related to the solution according to the present application are shown in the drawings, while other details less related to the present application are omitted.
[0034] It should be understood that the present application is not limited to the described embodiments due to the following description with reference to the drawings. In this document, where feasible, embodiments can be combined with each other, features can be replaced or borrowed between different embodiments, and one or more features can be omitted in one embodiment.
[0035] Sound pressure is a signal that can be measured in a non-contact manner in space. The pressure pulsations caused during the hot fire test of the engine cause the surrounding air medium to produce response pulsations. Therefore, monitoring the sound pressure pulsations of the engine can, to a certain extent, reflect the characteristics of the pressure pulsations. At the same time, its non-contact installation method facilitates its multi-point arrangement. In summary, the present application proposes a method for diagnosing the combustion instability excitation source of a rocket engine using an acoustic matrix, which has certain engineering application value.
[0036] Currently, when the entire engine undergoes a hot fire test, each component will generate its own distinct characteristic acoustic frequencies during engine operation, which are related to physical conditions such as the component structure size, temperature, and material. The acoustic frequencies of each component may also be transmitted from upstream to downstream through the connection structure or the propellant flow in the pipeline, and the sound waves formed when each component operates will "blend" together. During the propagation of sound waves, the pulsating changes in the sound field will also bring about pulsating changes in the pressure field, and this mechanism of action will become an excitation source for combustion instability under suitable conditions. When there are multiple excitation sources in the combustion chamber, correctly identifying the sources of different frequency excitation sources and avoiding frequency bands close to the acoustic frequencies of the combustion chamber is crucial for designers.
[0037] The embodiments of the present application provide a method for diagnosing the combustion instability excitation source of a rocket engine using an acoustic matrix, Figure 1 which shows a flow block diagram of the method for diagnosing the combustion instability excitation source of a rocket engine using an acoustic matrix. Refer to Figure 1 and the method includes:
[0038] Step S1, obtaining sound pressure signals monitored by a plurality of sound pressure sensors uniformly arranged along the axial and circumferential directions of the rocket engine.
[0039] Here, multiple sound pressure sensors are evenly arranged along the axial direction of the rocket engine, from the top of the rocket engine to the nozzle exit, that is, the rocket engine is measured in layers along the axial direction; at the same time, in the circumferential position, it is ensured that the adjacent span angles of the sound pressure sensors are the same. Based on the axial division of each layer, at least four sound pressure sensors are evenly distributed circumferentially. Figure 2 The schematic diagram of the arrangement of multiple sound pressure sensors along the axial direction of the rocket engine is shown. Figure 3 The schematic diagram of the arrangement of multiple sound pressure sensors along the circumferential direction of the rocket engine is shown.
[0040] Step S2: Process all sound pressure signals using independent component analysis (ICA) to obtain multiple sound pressure signals with different characteristic frequencies.
[0041] Step S3: Perform FFT (Fast Fourier Transform) on multiple sound pressure signals with different characteristic frequencies to determine the characteristic frequency of each sound pressure signal.
[0042] Step S4: Obtain the main frequency monitored in the engine combustion chamber, and determine the sound pressure signal corresponding to the characteristic frequency that is the same as the main frequency as the target sound pressure signal. Here, existing contact measurement methods, such as vibration pressure sensor and dynamic pressure sensor measurement methods, can be used to obtain the main frequency, that is, the frequency at which anomalies occur.
[0043] Step S5: The position where the sound pressure sensor corresponding to the target sound pressure signal is arranged is the position where the excitation source generated by the final determination is located.
[0044] In this embodiment, by arranging sound pressure sensors to collect the sound pressure signals during the hot test of the engine, and processing the collected signals through the independent component analysis method (ICA), the working frequencies of different components can be accurately obtained by back-calculation, so as to determine the excitation source caused by combustion instability; from an acoustic perspective, decouple the transfer relationship between each component to solve the problem of the influence of combustion instability of other components of the engine on the thrust chamber.
[0045] Based on the same inventive concept as the method for diagnosing the combustion instability excitation source using the acoustic matrix of the rocket engine, this embodiment also provides a corresponding device for diagnosing the combustion instability excitation source using the acoustic matrix of the rocket engine, including:
[0046] A sound pressure signal acquisition module for acquiring the sound pressure signals monitored by multiple sound pressure sensors evenly arranged along the axial and circumferential directions of the rocket engine;
[0047] A signal processing module for processing all sound pressure signals using independent component analysis to obtain multiple sound pressure signals with different characteristic frequencies;
[0048] A characteristic frequency determination module, configured to perform FFT on multiple sound pressure signals with different characteristic frequencies to determine the characteristic frequency of each sound pressure signal;
[0049] A target sound pressure signal determination module, configured to obtain the main frequency monitored in the engine combustion chamber and determine the sound pressure signal corresponding to the characteristic frequency that is the same as the main frequency as the target sound pressure signal;
[0050] An excitation source position determination module, configured to determine that the position where the sound pressure sensor corresponding to the target sound pressure signal is arranged is the position where the finally determined excitation source is generated.
[0051] The device for diagnosing the combustion instability excitation source of the rocket engine by the acoustic matrix in this embodiment has the same inventive concept as the method for diagnosing the combustion instability excitation source of the rocket engine by the acoustic matrix above. Therefore, the specific implementation manner of this device can be seen in the embodiment part of the method for diagnosing the combustion instability excitation source of the rocket engine by the acoustic matrix in the previous text, and its technical effects correspond to those of the above method, which will not be elaborated here.
[0052] This application embodiment also provides a system for diagnosing the combustion instability excitation source of the rocket engine by the acoustic matrix, including: a rocket engine 3, a data transmission line 1, multiple sound pressure sensors 2, a data receiver 4, and a data processor 5;
[0053] Multiple sound pressure sensors 2 are uniformly arranged along the axial and circumferential directions of the rocket engine 3 for detecting sound pressure signals at different positions;
[0054] The sound pressure signals at different positions are sent to the data receiver 4 through the data transmission line 1;
[0055] The data receiver 4 sends the sound pressure signals at different positions to the data processor 5; the data processor 5 is configured to implement the method for diagnosing the combustion instability excitation source of the rocket engine in the foregoing embodiment.
[0056] In summary, this application has the following technical effects:
[0057] 1. By arranging sound pressure sensors, this application collects the sound pressure signals during the hot test of the engine, orthogonally processes the collected signals through the independent component analysis method (ICA), and can accurately obtain the working frequencies of different components after back-projection, and back-project the excitation source generated by combustion instability; from an acoustic perspective, it decouples the transfer relationships between various components and solves the problem of the influence of combustion instability of other components of the engine on the thrust chamber.
[0058] 2. This application can be applied to the hot tests of engines with various different structural dimensions. When the structural dimensions of the engine are different, the positions of the sound pressure sensors can be adjusted according to the maximum outer diameter dimension of the engine.
[0059] 3. This application does not require surface mounting on the engine, avoiding the installation difficulties of traditional contact sensors under high temperature, high pressure, and large vibration conditions.
[0060] 4. This application will collect sound pressure signals in different axial and circumferential directions of the engine, completely capture the sound pressure signals of the entire sound field, with high measurement accuracy, and at the same time, the sound pressure amplitude range that the sensor can measure is relatively large.
[0061] 5. The number of sound pressure sensors arranged in this application can be flexibly adjusted. When the number of sound pressure sensors is sufficient, several sound pressure sensors can still meet the usage measurement requirements in case of failure, with a high usage fault tolerance rate.
[0062] As described above, the above are only various implementation manners of this application, but the protection scope of this application is not limited thereto. Any person skilled in the art within the technical scope disclosed in this application can easily think of changes or substitutions, which should all be covered within the protection scope of this application. Therefore, the protection scope of this application shall be subject to the protection scope of the claims.
Claims
1. A method for diagnosing the excitation source of combustion instability in a rocket engine by using an acoustic matrix, characterized in that Including: Obtain the sound pressure signals monitored by a plurality of sound pressure sensors uniformly arranged along the axial and circumferential directions of the rocket engine; Process all the sound pressure signals using independent component analysis to obtain a plurality of sound pressure signals with different characteristic frequencies; Perform FFT on the plurality of sound pressure signals with different characteristic frequencies to determine the characteristic frequency of each sound pressure signal; Obtain the main frequency monitored in the engine combustion chamber, and determine the sound pressure signal corresponding to the characteristic frequency that is the same as the main frequency as the target sound pressure signal; The arrangement position of the sound pressure sensor corresponding to the target sound pressure signal is the position where the finally determined excitation source is generated.
2. The method according to claim 1, characterized in that The plurality of sound pressure sensors are uniformly arranged along the axial direction of the rocket engine, including: Uniformly arranged from the top of the rocket engine to the nozzle exit.
3. A device for diagnosing the excitation source of combustion instability in a rocket engine acoustic matrix, characterized in that, Including: A sound pressure signal acquisition module for obtaining the sound pressure signals monitored by a plurality of sound pressure sensors uniformly arranged along the axial and circumferential directions of the rocket engine; A signal processing module for processing all the sound pressure signals using independent component analysis to obtain a plurality of sound pressure signals with different characteristic frequencies; A characteristic frequency determination module for performing FFT on the plurality of sound pressure signals with different characteristic frequencies to determine the characteristic frequency of each sound pressure signal; A target sound pressure signal determination module for obtaining the main frequency monitored in the engine combustion chamber and determining the sound pressure signal corresponding to the characteristic frequency that is the same as the main frequency as the target sound pressure signal; An excitation source position determination module for the arrangement position of the sound pressure sensor corresponding to the target sound pressure signal being the position where the finally determined excitation source is generated.
4. The device according to claim 3, characterized in that, The plurality of sound pressure sensors are uniformly arranged along the axial direction of the rocket engine, including: Uniformly arranged from the top of the rocket engine to the nozzle exit.
5. A system for diagnosing the excitation source of combustion instability in a rocket engine acoustic matrix, characterized in that, Including: A rocket engine (3), a data transmission line (1), a plurality of sound pressure sensors (2), a data receiver (4), and a data processor (5); The plurality of sound pressure sensors (2) are uniformly arranged along the axial and circumferential directions of the rocket engine (3) for detecting sound pressure signals at different positions; The sound pressure signals at different positions are sent to the data receiver (4) through the data transmission line (1); The data receiver (4) sends the sound pressure signals at different positions to the data processor (5); the data processor (5) is used to implement the method for diagnosing the excitation source of combustion instability of the rocket engine acoustic matrix according to claim 1 or 2.