CN-120626377-B - Diagnostic method and device for unstable combustion oscillation characteristics of rocket engine

CN120626377BCN 120626377 BCN120626377 BCN 120626377BCN-120626377-B

Abstract

The application provides a method and a device for diagnosing unstable combustion oscillation characteristics of a rocket engine, and relates to the field of rocket engine testing, wherein the method comprises the steps of obtaining component parameters of each component of a thrust chamber of the rocket engine to be analyzed; the method comprises the steps of determining the acoustic admittance of a nozzle outlet based on the nozzle parameters and the channel parameters, determining the mass flow pulsation of the outlet of each nozzle channel, processing the design parameters of an engine combustion chamber and the configured flame parameters, determining the heat release pulsation of single flame, solving the configured acoustic feature network based on the configured modal cutoff value, the mass flow pulsation and the heat release pulsation to obtain a plurality of actual main frequencies and the modal numbers corresponding to the actual main frequencies, and determining the modal types of the actual main frequencies. According to the application, the actual main frequency of the low-temperature engine during oscillation combustion under different working conditions can be predicted according to the injection pulsation of a single nozzle or a plurality of nozzles and the dynamic response characteristic of the heat release pulsation of combustion flame, and related parameter analysis is carried out.

Inventors

YANG LIJUN
NAN JIAQI
LI JINGXUAN

Assignees

北京航空航天大学

Dates

Publication Date: 20260505
Application Date: 20250716

Claims (10)

1. A method of diagnosing unstable combustion oscillation characteristics of a rocket engine, the method comprising: Acquiring component parameters of each component of a thrust chamber of a rocket engine to be analyzed, wherein the component parameters comprise engine combustion chamber design parameters, nozzle parameters and channel parameters, and the channel parameters comprise Mach numbers of propellant channels; Determining a nozzle outlet acoustic admittance based on the nozzle parameters and the channel parameters; Processing the nozzle outlet acoustic admittance and the nozzle parameters to determine the mass flow pulsation of the outlet of each nozzle channel; Processing the design parameters of the engine combustion chamber and the configured flame parameters to determine the heat release pulsation of the single flame; Solving a configured acoustic feature network based on a configured modal cutoff value, the mass flow pulsation and the heat release pulsation to obtain a plurality of actual main frequencies and modal numbers corresponding to the actual main frequencies, wherein the acoustic feature network is an equation based on Fourier transformation and multi-modal superposition for constructing oscillation propagation characteristics of all components of the rocket engine to be analyzed; And determining the mode category of each actual main frequency and the three-dimensional space sound field distribution of pressure oscillation in the combustion chamber of the rocket engine to be analyzed based on the configured frequency reference sequence and the mode number of each actual main frequency.
2. The method of claim 1, wherein the configuring of the frequency reference sequence comprises: determining a plurality of self-excitation frequencies based on the nozzle outlet acoustic admittance and a configured start-stop frequency range; Processing the design parameters of the engine combustion chamber and the configured start-stop frequency range based on the mode cut-off value, and determining a plurality of mode frequencies under a passive condition; and sequencing the multiple self-excitation frequencies and the multiple modal frequencies in order from small to large to obtain a frequency reference sequence.
3. The method of claim 1, wherein the expression of the mass flow pulsation is: Wherein, the For the mass flow pulsation to be a function of, Is the first The cross-sectional area of the outlet of each nozzle, The amplitude of the pressure oscillations at the burner injection panel for the nozzle outlets, As a function of the transfer of the nozzle, , For the acoustic admittance of the nozzle outlet, For the mach number of the propellant channels, Is the average sound velocity of the low-temperature propellant, Is a complex number of circle frequencies.
4. The method of claim 1, wherein the expression of the heat release pulsation is: Wherein, the The heat release pulsation of the flame unit area corresponding to the j-th nozzle, The time soaking release rate is the single flame unit area, As an average pressure of the combustion chamber, The amplitude of the pressure oscillations at the burner injection panel for the nozzle outlets, As a transfer function of the j-th flame, Is a complex number of circle frequencies.
5. The method of claim 1, wherein the expression for the network of acoustic features is: Wherein, the Is dependent on complex circular frequency Is used for the coefficient matrix of (a), As the magnitude vector of the magnitude vector, Is a source item vector.
6. The method of claim 5, wherein solving the configured acoustic signature network based on the configured modal cutoff, the mass flow pulsation, and the heat release pulsation to obtain a plurality of actual dominant frequencies and a number of modes corresponding to each actual dominant frequency comprises: solving a configured acoustic feature network based on a configured modal cutoff value, the mass flow pulsation and the heat release pulsation to obtain a first number of feature complex frequencies; Determining an actual main frequency corresponding to each characteristic complex frequency based on the first number of characteristic complex frequencies; Substituting the characteristic complex frequency into a coefficient matrix of the acoustic characteristic network for any characteristic complex frequency to perform dimension reduction processing to obtain a dimension reduction coefficient matrix; And determining the mode number corresponding to each actual main frequency based on the maximum value of the reduced-dimension coefficient matrix.
7. The method of claim 6, wherein after determining the actual dominant frequency for each characteristic complex frequency based on the first number of characteristic complex frequencies, the method further comprises: increasing the modal cutoff value to obtain a current modal cutoff value; Solving the configured acoustic feature network based on the current modal cutoff value, the mass flow pulsation and the heat release pulsation to obtain a second number of actual main frequencies; if the first number is the same as the second number, determining that the number of the actual main frequencies meets the configured number condition, and returning to the execution step, wherein the characteristic complex frequency is substituted into a coefficient matrix of the acoustic characteristic network for any characteristic complex frequency to perform dimension reduction processing to obtain a dimension reduction coefficient matrix; if the first number is different from the second number, returning to the executing step, namely increasing the modal cutoff value to obtain the current modal cutoff value.
8. A diagnostic device for unstable combustion oscillation characteristics of a rocket engine, the device comprising: the acquisition unit is used for acquiring component parameters of each component of a thrust chamber of the rocket engine to be analyzed, wherein the component parameters comprise engine combustion chamber design parameters, nozzle parameters and channel parameters; a determining unit that determines a nozzle outlet acoustic admittance based on the nozzle parameter and the channel parameter; Processing the nozzle outlet acoustic admittance and the nozzle parameters to determine the mass flow pulsation of the outlet of each nozzle channel; processing the design parameters of the engine combustion chamber and the configured flame parameters to determine the heat release pulsation of the single flame; the system comprises a calculation unit, a configuration unit and a heat release unit, wherein the calculation unit is used for solving a configuration acoustic feature network based on a configuration modal cut-off value, the mass flow pulsation and the heat release pulsation to obtain a plurality of actual main frequencies and modal numbers corresponding to the actual main frequencies; the determining unit is used for determining the mode category of each actual main frequency and the three-dimensional space sound field distribution of pressure oscillation in the combustion chamber of the rocket engine to be analyzed based on the configured frequency reference sequence and the mode number of each actual main frequency.
9. An electronic device, characterized in that the electronic device comprises a processor, a communication interface, a memory and a communication bus, wherein the processor, the communication interface and the memory are in communication with each other through the communication bus; a memory for storing a computer program; a processor for executing a computer program stored on a memory for performing the method steps of any of claims 1-7.
10. A computer-readable storage medium, characterized in that the computer-readable storage medium has stored therein a computer program which, when executed by a processor, implements the method steps of any of claims 1-7.

Description

Diagnostic method and device for unstable combustion oscillation characteristics of rocket engine Technical Field The application relates to the field of rocket engine testing, in particular to a method and a device for diagnosing unstable combustion oscillation characteristics of a rocket engine. Background The liquid rocket engine is an important power device for the carrier rocket and the strategic weapon, and the low-temperature liquid rocket engine is becoming the first choice of the high-thrust repeatable carrier rocket power due to the advantages of high performance, high cleanliness, no toxicity, environmental protection and the like. In the design process of a core component, namely a thrust chamber, unlike a conventional propellant, the low-temperature propellant is often in a trans/supercritical state (with compressibility similar to gas and high density similar to liquid) under high-temperature and high-pressure working conditions, the injection process is more susceptible to the pressure oscillation of a combustion chamber, so that flow pulsation and combustion heat release pulsation with larger amplitude and higher frequency are generated, and when instability occurs, the pressure oscillation in the combustion chamber often causes the phenomenon that the dominant frequency of the injection pulsation of a nozzle (called a nozzle coupling mode) and the natural acoustic frequency of the combustion chamber (called a combustion chamber acoustic mode) coexist. The unstable combustion of nozzle coupling is often not controlled or eliminated effectively by the traditional suppression devices such as an acoustic cavity, so that the problem of unstable combustion of nozzle coupling of a low-temperature engine is solved, once the unstable combustion occurs, the engine vibrates and shortens the service life, and the explosion is caused by heavy weight. However, the cost of the optimal design is extremely high through a full-size engine test, and the period of the coupling multi-physical-field process such as injection, combustion, pressure oscillation and the like with huge space-time scale difference is extremely long by adopting a full-size high-precision numerical method, so that the method is difficult to be directly used for the optimal design of a wide-working-condition range real low-temperature liquid rocket engine. The development process of the high-thrust repeatable low-temperature liquid rocket engine such as liquid hydrogen, liquid oxygen methane and the like is greatly limited. Disclosure of Invention The embodiment of the application aims to provide a method and a device for diagnosing unstable combustion oscillation characteristics of a rocket engine, which are used for solving the technical problems of rapid diagnosis and accurate regulation of unstable combustion oscillation frequency types lacking nozzle coupling in low-temperature liquid rocket engine design and test run technology. In a first aspect, a method for diagnosing unstable combustion oscillation characteristics of a rocket engine is provided, which may include: Acquiring component parameters of each component of a thrust chamber of a rocket engine to be analyzed, wherein the component parameters comprise engine combustion chamber design parameters, nozzle parameters and channel parameters; Determining a nozzle outlet acoustic admittance based on the nozzle parameters and the channel parameters; Processing the nozzle outlet acoustic admittance and the nozzle parameters to determine the mass flow pulsation of the outlet of each nozzle channel; Processing the design parameters of the engine combustion chamber and the configured flame parameters to determine the heat release pulsation of the single flame; Solving a configured acoustic feature network based on a configured modal cutoff value, the mass flow pulsation and the heat release pulsation to obtain a plurality of actual main frequencies and modal numbers corresponding to the actual main frequencies, wherein the acoustic feature network is an equation based on Fourier transformation and multi-modal superposition for constructing oscillation propagation characteristics of all components of the rocket engine to be analyzed; And determining the mode category of each actual main frequency and the three-dimensional space sound field distribution of pressure oscillation in the combustion chamber of the rocket engine to be analyzed based on the configured frequency reference sequence and the mode number of each actual main frequency. In one possible implementation, the configuring procedure of the frequency reference sequence includes: determining a plurality of self-excitation frequencies based on the nozzle outlet acoustic admittance and a configured start-stop frequency range; Processing the design parameters of the engine combustion chamber and the configured start-stop frequency range based on the mode cut-off value, and determining a plurality of mode frequencies under a pa