CN112664355B - Method and device for measuring combustion speed of propellant of solid rocket engine - Google Patents
Method and device for measuring combustion speed of propellant of solid rocket engine Download PDFInfo
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Abstract
The invention discloses a method and a device for measuring the burning rate of a propellant of a solid rocket engine. When the axial burning rate is measured, the intermediate frequency signal obtained by mixing the incident signal and the reflected signal is analyzed by using the Doppler effect, so that the axial burning rate of the propellant is calculated, the round-trip time of electromagnetic waves in the propellant is not considered, the measurement error is reduced to a great extent, and the measurement precision is improved; and secondly, the axial burning rate and the transverse burning rate of the solid propellant can be measured simultaneously, and when the transverse burning rate of the propellant is measured, the tomography is combined to perform inversion imaging on the internal combustion condition of the rocket engine to obtain the transverse pushing distance of the propellant, so that the transverse burning rate of the propellant is obtained through calculation, the aim of multilayer detection is fulfilled, the cost can be reduced, and the dynamic, quick and accurate measurement of the burning rate of the propellant of the solid rocket engine is realized.
Description
Technical Field
The invention belongs to the field of solid rocket engine measurement, and particularly relates to a method, a device, a computing platform and a computer readable storage medium for measuring the combustion speed of a propellant of a solid rocket engine.
Background
The solid rocket engine technology is the core technology of space transportation and missile weapons. The solid propellant combustion characteristic influences the flight speed and range of the rocket, and the combustion speed of the combustion surface propulsion is a core parameter for representing the ballistic performance in an engine and is a key index needing strict control in the propellant formula design and production. Therefore, the calculation of the transverse burning rate and the axial burning rate of the solid propellant has important significance for discovering the essential law of propellant combustion, improving the reliability and the safety of solid rocket propellant combustion and meeting the ever-increasing application requirements of aerospace and military power in China.
The current mature burning rate test means include static test of a chemical strip, dynamic test of an engine, and test of rotary overload combustion and pressure transient combustion. Wherein, the target line method and the acoustic emission method can only test the average burning rate under specific pressure; the optical camera dynamic burning rate test is influenced by smoke deposition during propellant burning, nitrogen flow is required to be filled, X-ray imaging equipment is huge and radioactive, the ultrasonic penetration capacity is limited, and the measurement time is long. The methods have low time resolution, are difficult to realize dynamic real-time measurement of a combustion area, cannot realize comprehensive reappearance of the whole combustion process, cannot realize multi-level detection of the combustion rate of the propellant, are relatively complex in calculation process and relatively high in overall equipment cost, and have relatively large errors in the calculated combustion rate of the propellant.
Disclosure of Invention
In view of the above, the invention aims to realize dynamic, rapid and accurate measurement of the combustion speed of the propellant of the solid rocket engine.
In order to achieve the aim, the invention provides a solid rocket engine propellant burning rate measuring device, which comprises a detection array unit, a detection array control unit, a signal transmitting unit, a signal receiving unit, a burning rate measuring unit and a measuring time sequence control unit; wherein the content of the first and second substances,
the detection array unit is used for arranging detection nodes around the solid rocket engine, the detection nodes are distributed on M planes which are vertical to the axial direction of the solid rocket engine, N detection nodes are arranged on each plane, the detection array unit sends modulation signals to the solid rocket engine and receives signals after the modulation signals are reflected by the fuel surface of a propellant of the solid rocket engine and signals after the modulation signals are attenuated by the solid rocket engine, wherein M is more than or equal to 2, and N is more than or equal to 2;
the detection array control unit controls the connection of a detection node in the signal detection array unit and the signal transmitting unit or the connection of the signal receiving unit, the connection of the detection node and the signal transmitting unit is used as a transmitting node, and the connection of the detection node and the signal receiving unit is used as a receiving node;
the signal transmitting unit generates a modulation signal, and the modulation signal is transmitted to the transmitting node of the detection array unit and is transmitted to the burning rate measuring unit for measurement;
the signal receiving unit is used for receiving signals received by the receiving nodes of the detection array unit and sending the signals to the burning rate measuring unit for measurement;
the combustion speed measuring unit at least comprises an axial combustion speed measuring unit, and the axial combustion speed measuring unit analyzes an intermediate frequency signal obtained by mixing an incident signal transmitted by the transmitting node and a reflected signal received by the receiving node by using a Doppler effect so as to calculate the axial combustion speed of the propellant;
and the measurement time sequence control unit is used for obtaining the working time sequences of the units according to the measurement task requirements and continuously measuring the burning rate of the propellant of the solid rocket engine, wherein the burning rate at least comprises the axial burning rate.
Furthermore, the burning rate measuring device also comprises a transverse burning rate measuring unit, the transverse burning rate measuring unit analyzes the signal after the modulation signal sent by the signal receiving unit is attenuated by the solid rocket engine, and inversion imaging is carried out on the propellant in the rocket engine by utilizing a tomography algorithm to obtain the transverse pushing distance of the propellant, so that the transverse burning rate of the propellant is obtained.
The invention also provides a method for measuring the burning rate of the propellant of the solid rocket engine, which adopts the device and comprises the following steps:
s1, before combustion t0The step of measuring the solid propellant at the moment, in which,
a pair of detection nodes positioned on different detection planes form a transceiving pair, before the solid propellant burns, one detection node sends a modulation signal with the center frequency f, the other detection node receives a signal of the modulation signal reflected by the end face of the solid propellant, the transmission signal and the reception signal are mixed to obtain a first intermediate frequency signal, further the frequency delta f of the first intermediate frequency is obtained, and the initial end face distance L of the solid propellant is obtained through calculation0，
Wherein, B represents the modulation bandwidth of the modulation signal, T is the modulation period of the modulation signal, and c is the speed of light;
s2, a step of measurement of the solid propellant at the time of combustion, in which step,
at tnAt the moment, N is 1,2, … and N, a pair of detection nodes located on different detection planes form a transceiving pair, when the solid propellant burns, one detection node sends a modulation signal with the center frequency of f, and the other detection node receives the modulation signal reflected by the end face of the solid propellantThe signal, the transmitting signal and the receiving signal are mixed to obtain a second intermediate frequency signal, and then the frequency of the second intermediate frequency signal is obtained
And
indicating the frequency of the output signal in the forward modulation band after the frequency mixing processing of the transmitting signal and the receiving signal,
representing the frequency of the output signal of the transmitting signal and the receiving signal in a negative modulation frequency band after frequency mixing processing, and then calculating the solid propellant tnAxial burning velocity V of timenAnd end face distance Ln：
Further, in step S2, the axial burning velocity V is measurednAnd end face distance LnThe receiving and transmitting pairs have multiple groups, and the axial burning velocity V is obtained by using the transmitting signals and the receiving signals of the multiple groups of receiving and transmitting pairsnAnd end face distance LnStatistical averaging is performed to reduce errors.
Further, in step S2, at tnThe following measurements are simultaneously carried out at the moment, detection nodes positioned on the same detection plane form a transceiving pair, when the solid propellant burns, one detection node sends a modulation signal with the center frequency of f, other detection nodes receive the signal of which the modulation signal is projected through the solid rocket engine, and t is obtained through tomographynAt the moment of propellant thickness r on the detection planen(ii) a In addition, also comprises
Step S3, measurement of propellant thickness in the above step S2 at N moments, based on propellant initiationThickness r0And the thickness r of a certain detection plane at N moments measured in the step S2nSet of compositions rnN is 0,1,2, …, N }, and the solid propellant can be obtained at any two times tiAnd tjTransverse burning rate rij，
The invention also provides a computer readable storage medium, which stores a computer program, and is characterized in that the computer program is executed by a processor to realize the method for measuring the burning rate of the solid rocket engine propellant.
The invention also provides a solid rocket engine propellant burning rate measuring and calculating platform, which is characterized by comprising the following components:
at least one processor; and (c) a second step of,
a memory communicatively coupled to the at least one processor; wherein, the first and the second end of the pipe are connected with each other,
the memory stores instructions executable by the at least one processor to enable the at least one processor to perform the solid rocket engine propellant burning rate measurement method described above.
Advantageous effects
When the axial burning rate is measured, the Doppler effect is utilized to analyze an intermediate frequency signal obtained after mixing an incident signal and a reflected signal, so that the axial burning rate of the propellant is calculated, the round-trip time of electromagnetic waves in the propellant is not considered, the measurement error is reduced to a great extent, and the measurement precision is improved; and secondly, the axial burning rate and the transverse burning rate of the solid propellant can be measured simultaneously, and when the transverse burning rate of the propellant is measured, the tomography is combined to perform inversion imaging on the internal combustion condition of the rocket engine to obtain the transverse pushing distance of the propellant, so that the transverse burning rate of the propellant is obtained through calculation, the aim of multilayer detection is fulfilled, the cost can be reduced, and the dynamic, quick and accurate measurement of the burning rate of the propellant of the solid rocket engine is realized.
Drawings
FIG. 1 is a block diagram of the system components of the propellant burning rate measuring device of the solid rocket engine of the present invention;
FIG. 2 is a schematic diagram of the arrangement positions of the probing nodes according to the present invention;
FIG. 3 is a schematic view of the principle of the measurement of the propellant burning rate of the solid rocket engine of the present invention;
FIG. 4 is a schematic representation of the frequency of the transmitted and received signals as a function of time for a measurement of the unburned propellant in a solid rocket engine of the present invention;
FIG. 5 is a schematic diagram showing the variation of the frequency of an intermediate frequency signal with time after mixing of a transmitted signal and a received signal during measurement of unburned propellant in a solid rocket engine according to the present invention;
FIG. 6 is a schematic diagram of the frequency of the transmitted signal, the received signal, and the intermediate frequency signal as a function of time for a solid rocket engine propellant combustion measurement in accordance with the present invention;
FIG. 7 is a schematic diagram showing the variation of the frequency of the intermediate frequency signal with time obtained after difference frequency is carried out on the upper frequency sweep and the lower frequency sweep of the transmitting signal and the receiving signal during the propellant combustion measurement of the solid rocket engine.
Detailed Description
The following describes in detail embodiments of the present invention with reference to the drawings.
As shown in figure 1, the solid rocket engine propellant burning rate measuring device comprises a detection array unit, a detection array control unit, a signal transmitting unit, a signal receiving unit, a burning rate measuring unit and a measuring time sequence control unit; the detection array unit is used for arranging detection nodes around the solid rocket engine, the detection nodes are distributed on M planes which are vertical to the axial direction of the solid rocket engine, N detection nodes are arranged on each plane, the detection array unit sends modulation signals to the solid rocket engine and receives signals after the modulation signals are reflected by a propellant combustion surface of the solid rocket engine and signals after the modulation signals are attenuated after the modulation signals are transmitted through the solid rocket engine; the detection array control unit controls the connection of a detection node in the signal detection array unit and the signal transmitting unit or the connection of the signal receiving unit, the connection of the detection node and the signal transmitting unit is used as a transmitting node, and the connection of the detection node and the signal receiving unit is used as a receiving node; the signal transmitting unit generates a modulation signal, and the modulation signal is transmitted to the transmitting node of the detection array unit and is transmitted to the burning rate measuring unit for measurement; the signal receiving unit is used for receiving signals received by the receiving nodes of the detection array unit and sending the signals to the burning rate measuring unit for measurement; the combustion rate measuring unit at least comprises an axial combustion rate measuring unit, and the axial combustion rate measuring unit analyzes an intermediate frequency signal obtained by mixing an incident signal transmitted by the transmitting node and a reflected signal received by the receiving node by using a Doppler effect so as to calculate the axial combustion rate of the propellant; and the measurement time sequence control unit is used for obtaining the working time sequences of the units according to the measurement task requirements and continuously measuring the burning rate of the propellant of the solid rocket engine, wherein the burning rate at least comprises the axial burning rate.
Furthermore, the burning rate measuring device also comprises a transverse burning rate measuring unit, the transverse burning rate measuring unit analyzes the signal after the modulation signal sent by the signal receiving unit is attenuated by the solid rocket engine, and inversion imaging is carried out on the propellant in the rocket engine by utilizing a tomography algorithm to obtain the transverse pushing distance of the propellant, so that the transverse burning rate of the propellant is obtained.
The detection nodes of the embodiment of the invention are arranged on a plurality of planes vertical to the axial direction of the solid rocket engine, as shown in fig. 2, the detection nodes of the embodiment are arranged on a plurality of circumferences vertical to the axial direction of the solid rocket engine, and can also be arranged in other ways on the same plane. Each detection node has a transceiving function so as to achieve the purpose of multilayer detection of the burning rate of the propellant, the detection nodes which are not on the same plane can receive reflected signals of waves burning at different stages, frequency mixing processing is carried out on the reflected signals and incident signals to measure Doppler frequency, the axial burning rate of the propellant is calculated by utilizing the Doppler frequency, attenuation information of the waves in the solid propellant is obtained in the transceiving nodes on the same plane, the back imaging is carried out on the burning surface of the solid propellant by combining a tomography technology, the visual presentation of the transverse diffusion distance of the propellant in a combustion chamber is realized, and the transverse burning rate of the solid propellant is further calculated.
The method for measuring the burning rate of the solid rocket engine propellant utilizes the following basic principles:
(1) principle for measuring axial burning rate
When the solid propellant burns, the propellant burning surface is axially pushed, as shown in fig. 2 and fig. 3, the transceiving node 1 transmits an electromagnetic wave with frequency f, the transceiving nodes 2 and/or 5 receive the electromagnetic wave which is transmitted by the transceiving node 1 and reflected by the propellant burning surface, or the transceiving node 3 transmits an electromagnetic wave with frequency f, the transceiving nodes 4 and/or 6 receive the electromagnetic wave which is transmitted by the transceiving node 1 and reflected by the propellant burning surface, and the received electromagnetic wave has frequency:
wherein c is the propagation velocity of electromagnetic wave, v0The axial combustion speed of the solid propellant is shown, the combustion surface in the figure is pushed from left to right, when the combustion surface approaches to the detection node 1 or 3, the operation symbol in the formula is a plus sign, and when the combustion surface is far away from the detection node 1 or 3, the operation symbol in the formula is a minus sign.
The doppler frequency is then:
when the receiving and transmitting nodes 2, 4, 5, 6 receive the reflected signals, a doppler effect will occur, which generates a doppler frequency, and the total transmitting and receiving process is equivalent to two doppler effects, therefore, the doppler frequency fdComprises the following steps:
from the above formula, one can obtain:
therefore, as long as the Doppler frequency is measured, the axial burning rate of the propellant can be calculated.
When the propellant is not burnt, a pair of detection nodes located on different circumferences form a transceiving pair, for example, 1 and 2 (or 1 and 5, 3 and 4, 3 and 6), and the change law of the frequency of the transmission signal and the frequency of the receiving signal along with time is measured as shown in fig. 4.
There is a time difference at between the transmitted signal and the received signal,
in the formula, R represents the initial distance of the propellant combustion surface, the transmitting signal and the receiving signal are mixed to obtain an intermediate frequency signal, and the change curve of the frequency delta f of the intermediate frequency signal along with time is shown in figure 5.
From the graph geometry in fig. 4, it can be derived:
where B represents the modulation bandwidth and T is the modulation period. Combining the formula 2 and the formula 3, the initial distance L of the end face of the solid propellant when the solid propellant is not combusted can be obtained0：
As can be seen from the above formula, the distance L of the end face of the solid propellant can be obtained by only obtaining the frequency delta f of the intermediate frequency signal0The intermediate frequency signal frequency can be obtained by spectral analysis.
Doppler effect is caused by axial displacement generated by propellant combustion, and Doppler frequency shift f is mixed in signals received by the transmitting and receiving nodesdThe frequency of the transmitted and received signals varying with time andthe frequency signal frequency variation is shown in fig. 6.
In FIG. 6, fb+And fb-The frequencies of the intermediate frequency signal after difference frequency is respectively carried out on the upper frequency sweep and the lower frequency sweep of the transmitting signal and the receiving signal, and fdThe relationship of (1) is:
fb+＝Δf-fdequation 5
fb-＝Δf+fdEquation 6
Where Δ f is the frequency of the intermediate frequency signal when the propellant is burning, which can be obtained in step S1.
The axial burning velocity V of the propellant can be obtained by simultaneous formulas 1, 5 and 6:
the axial displacement L of the propellant can be found by the simultaneous equations 4, 5 and 6:
where f is the intermediate frequency signal center frequency.
(2) Measuring principle of transverse burning rate
When the solid propellant burns, the propellant burning surface can be transversely pushed, as shown in fig. 2 and fig. 3, at this time, the transmitting and receiving nodes 1,2 and 5 transmit electromagnetic waves, and the transmitting and receiving nodes 3, 4 and 6 in the same plane correspondingly receive transmission signals. In the combustion process, when the detection node is used for carrying out first measurement, the meat thickness r of the propellant is obtained through inversion by combining the tomography technology1And when the node is detected for the second measurement, the thickness r of the propellant is obtained2And the time interval of the two measurements is delta t, the transverse burning rate of the propellant is as follows:
when propellant is burnt, chromatography is carried out by using antenna arrayImaging, inverting the transverse diffusion condition of the combustion surface of the propellant, and obtaining the thickness r of the propellant when detecting the node for the first measurement in the combustion process1And when the node is detected for the second measurement, the thickness r of the propellant is obtained2The time interval of the two measurements is delta t; from the measured data, the cross-fire rate of the solid propellant can be calculated as:
based on the basic principle, the method for measuring the burning rate of the propellant of the solid rocket engine adopts the measuring device, the detection node is placed 100mm away from the shell, and the method comprises the following steps:
s1, pre-combustion t0The step of measuring the solid propellant at the moment, in which,
a pair of detection nodes positioned on different detection planes form a transceiving pair, before the solid propellant burns, one detection node sends a modulation signal with the center frequency f, the other detection node receives a signal of the modulation signal reflected by the end face of the solid propellant, the transmission signal and the reception signal are mixed to obtain a first intermediate frequency signal, further the frequency delta f of the first intermediate frequency is obtained, and the initial end face distance L of the solid propellant is obtained through calculation0，
Wherein, B represents the modulation bandwidth of the modulation signal, T is the modulation period of the modulation signal, and c is the speed of light;
s2, a step of measurement of the solid propellant at the time of combustion, in which step,
at tnAt time, N is 1,2, … and N, a pair of detection nodes located on different detection planes form a transceiving pair, when the solid propellant burns, one detection node sends a modulation signal with the center frequency f, and the other detection node receives a signal of the modulation signal reflected by the end face of the solid propellant, and transmits and receives the signalAfter the signals are mixed, a second intermediate frequency signal is obtained, and then the frequency of the second intermediate frequency signal is obtained
And
indicating the frequency of the output signal in the forward modulation band after the frequency mixing processing of the transmitting signal and the receiving signal,
representing the frequency of the output signal of the transmitting signal and the receiving signal in a negative modulation frequency band after frequency mixing processing, and then calculating the solid propellant tnAxial burning velocity V of timenAnd end face distance Ln：
At tnThe following measurements are simultaneously carried out at the moment, detection nodes on the same detection plane form a transceiving pair, when the solid propellant burns, one detection node sends a modulation signal with the center frequency of f, other detection nodes receive the signal of the modulation signal which is projected by the solid rocket engine, and t is obtained through tomographynAt the moment of time the propellant thickness r on the circumferencen
S3, measuring the propellant thickness in the step S2 at N moments according to the initial propellant thickness r0And the thickness r of a certain circle at N moments measured in the step S2nSet of compositions rnN is 0,1,2, …, N }, and the solid propellant can be obtained at any two times tiAnd tjTransverse burning rate rij，
The burning rate of a certain section of the solid propellant is determined, and as a plurality of detection nodes are arranged around the rocket engine, the axial burning rate and the transverse burning rate of the related nodes are calculated according to the method, and finally the average value is obtained, so that the obtained result is more accurate. Because 1% of combustion speed error can bring up to 7% of thrust error, therefore improve the measurement accuracy of solid propellant combustion speed, have important meaning to successfully designing the solid rocket engine.
According to the invention, when the axial burning rate is measured, the Doppler effect is utilized to analyze the intermediate frequency signal obtained after the incident signal and the reflected signal are mixed, so that the axial burning rate of the propellant is calculated, the round-trip time of electromagnetic waves in the propellant is not considered, the measurement error is reduced to a great extent, the measurement accuracy is improved, and meanwhile, when the transverse burning rate of the propellant is measured, the internal combustion condition of the rocket engine is subjected to inversion imaging in combination with tomography, so that the transverse moving distance of the propellant is obtained, and the transverse burning rate of the propellant is calculated. The prior art can only measure the transverse burning rate and the axial burning rate of the propellant singly, and cannot achieve the aim of multilayer detection, which is an advantage that the prior art does not have.
The second embodiment of the invention relates to a solid rocket engine propellant burning rate measuring and calculating platform, which comprises: at least one processor; and a memory communicatively coupled to the at least one processor; wherein the memory stores instructions executable by the at least one processor to enable the at least one processor to perform the above-described method embodiments.
Where the memory and processor are connected by a bus, the bus may comprise any number of interconnected buses and bridges, the buses connecting together one or more of the various circuits of the processor and the memory. The bus may also connect various other circuits such as peripherals, voltage regulators, power management circuits, and the like, which are well known in the art, and therefore, will not be described any further herein. A bus interface provides an interface between the bus and the transceiver. The transceiver may be one element or a plurality of elements, such as a plurality of receivers and transmitters, providing a means for communicating with various other apparatus over a transmission medium. The data processed by the processor is transmitted over a wireless medium via an antenna, which further receives the data and transmits the data to the processor.
The processor is responsible for managing the bus and general processing and may also provide various functions including timing, peripheral interfaces, voltage regulation, power management, and other control functions. And the memory may be used to store data used by the processor in performing operations.
A third embodiment of the present invention relates to a computer-readable storage medium storing a computer program. The computer program realizes the embodiments of the method described above when executed by a processor.
That is, as can be understood by those skilled in the art, all or part of the steps in the method for implementing the embodiments described above may be implemented by a program instructing related hardware, where the program is stored in a storage medium and includes several instructions to enable a device (which may be a single chip, a chip, or the like) or a processor (processor) to execute all or part of the steps of the method described in the embodiments of the present application. And the aforementioned storage medium includes: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a magnetic disk, or an optical disk, and various media capable of storing program codes.
The present invention is not limited to the above preferred embodiments, and any modifications, equivalents, improvements, etc. within the principle of the idea of the present invention should be included in the protection scope of the present invention.
Claims (7)
1. A solid rocket engine propellant burning rate measuring device is characterized by comprising a detection array unit, a detection array control unit, a signal transmitting unit, a signal receiving unit, a burning rate measuring unit and a measuring time sequence control unit; wherein the content of the first and second substances,
the detection array unit is used for arranging detection nodes around the solid rocket engine, the detection nodes are distributed on M planes which are vertical to the axial direction of the solid rocket engine, N detection nodes are arranged on each plane, the detection array unit sends modulation signals to the solid rocket engine and receives signals after the modulation signals are reflected by the fuel surface of a propellant of the solid rocket engine and signals after the modulation signals are attenuated by the solid rocket engine, wherein M is more than or equal to 2, and N is more than or equal to 2;
the detection array control unit controls the connection of the detection nodes in the detection array unit and the signal transmitting unit or the connection of the signal receiving unit, the connection of the detection nodes and the signal transmitting unit is used as transmitting nodes, and the connection of the detection nodes and the signal receiving unit is used as receiving nodes;
the signal transmitting unit generates a modulation signal, and the modulation signal is transmitted to the transmitting node of the detection array unit and is transmitted to the burning rate measuring unit for measurement;
the signal receiving unit is used for receiving signals received by the receiving nodes of the detection array unit and sending the signals to the burning rate measuring unit for measurement;
the combustion rate measuring unit at least comprises an axial combustion rate measuring unit, and the axial combustion rate measuring unit analyzes an intermediate frequency signal obtained by mixing an incident signal transmitted by the transmitting node and a reflected signal received by the receiving node by using a Doppler effect so as to calculate the axial combustion rate of the propellant;
and the measurement time sequence control unit is used for obtaining the working time sequences of all the units according to the measurement task requirements and continuously measuring the burning rate of the propellant of the solid rocket engine, wherein the burning rate at least comprises the axial burning rate.
2. The solid rocket engine propellant burning rate measuring device of claim 1, wherein the burning rate measuring device further comprises a transverse burning rate measuring unit, the transverse burning rate measuring unit analyzes the signal after the modulation signal sent by the signal receiving unit is transmitted through the solid rocket engine and attenuated, and utilizes a tomography algorithm to carry out inversion imaging on the propellant in the rocket engine to obtain the transverse moving distance of the propellant, thereby obtaining the transverse burning rate of the propellant.
3. A method for measuring the burning rate of a propellant of a solid rocket engine, which adopts the device of claim 1 or 2, and is characterized by comprising the following steps:
s1, pre-combustion t0The step of measuring the solid propellant at the moment, in which,
a pair of detection nodes positioned on different detection planes form a transceiving pair, before the solid propellant burns, one detection node sends a modulation signal with the center frequency f, the other detection node receives a signal of the modulation signal reflected by the end face of the solid propellant, the transmission signal and the reception signal are mixed to obtain a first intermediate frequency signal, further the frequency delta f of the first intermediate frequency is obtained, and the initial end face distance L of the solid propellant is obtained through calculation0，
Wherein, B represents the modulation bandwidth of the modulation signal, T is the modulation period of the modulation signal, and c is the speed of light;
s2, step of measurement of the solid propellant at the time of combustion, in which step,
at tnAt the moment, N is 1,2, …, N, a pair of detection nodes located on different detection planes form a transceiving pair, when the solid propellant burns, one detection node sends a modulation signal with the center frequency of f, the other detection node receives a signal of the modulation signal reflected by the end face of the solid propellant, a second intermediate frequency signal is obtained after the frequency mixing of the sending signal and the receiving signal, and the frequency of the second intermediate frequency signal is further obtained
And
indicating the frequency of the output signal in the forward modulation band after the frequency mixing processing of the transmitting signal and the receiving signal,
representing the frequency of the output signal of the transmitting signal and the receiving signal in a negative modulation frequency band after frequency mixing processing, and then calculating the solid propellant tnAxial burning velocity V of timenAnd end face distance Ln：
4. A solid rocket engine propellant burning rate measuring method as claimed in claim 3, characterized in that in step S2, the axial burning rate V is measurednAnd end face distance LnThe receiving and transmitting pairs have multiple groups, and the axial burning velocity V is obtained by using the transmitting signals and the receiving signals of the multiple groups of receiving and transmitting pairsnAnd end face distance LnStatistical averaging is performed to reduce errors.
5. A solid rocket engine propellant burning rate measuring method as claimed in claim 3 or 4, characterized in that in step S2, at tnThe following measurements are simultaneously carried out at the moment, detection nodes on the same detection plane form a transceiving pair, when the solid propellant burns, one detection node sends a modulation signal with the center frequency of f, other detection nodes receive the signal of the modulation signal which is projected by the solid rocket engine, and t is obtained through tomographynAt the moment of propellant thickness r on the detection planen(ii) a In addition, also includes
Step S3, measuring propellant thickness in the step S2 at N moments according to propellant initial thickness r0And the thickness r of a certain detection plane at N moments measured in the step S2nSet of compositions rnN is 0,1,2, …, N }, and the solid propellant can be obtained at any two times tiAnd tjTransverse burning rate rij，
6. A computer-readable storage medium, storing a computer program, wherein the computer program, when executed by a processor, implements the method of measuring the propellant burning rate of a solid rocket engine according to any one of claims 3 to 5.
7. A solid rocket engine propellant burning rate measuring and calculating platform is characterized by comprising:
at least one processor; and the number of the first and second groups,
a memory communicatively coupled to the at least one processor; wherein the content of the first and second substances,
the memory stores instructions executable by the at least one processor to cause the at least one processor to perform the method of measuring solid rocket engine propellant burning rate of any one of claims 3-5.
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| RU2177113C1 (en) * | 2000-05-15 | 2001-12-20 | Федеральное государственное унитарное предприятие "Пермский завод им. С.М. Кирова" | Device for measurement of propellant burning rate in solid-propellant rocket engine |
| JP2010236425A (en) * | 2009-03-31 | 2010-10-21 | Nof Corp | Combustion speed measurement device, and measurement method using the same |
| RU2406864C1 (en) * | 2009-09-02 | 2010-12-20 | Федеральное государственное унитарное предприятие "Федеральный центр двойных технологий "Союз" (ФГУП "ФЦДТ "Союз") | Plant for determining burning speed of solid rocket fuel |
| CN104345118B (en) * | 2013-07-29 | 2016-08-10 | 西安电子科技大学 | Solid propellant many targets line Dynamic Burning Performance Test System and method |
| CN106353449B (en) * | 2016-11-03 | 2019-04-30 | 上海理工大学 | Active laser formula solid rocket propellant burn rate dynamic checkout unit and method |
| CN106482790B (en) * | 2016-11-09 | 2018-10-12 | 四川航天机电工程研究所 | Solid rocket propellant combustion measurement device based on Fire Radiation and measurement method |
| CN108167090B (en) * | 2018-02-08 | 2019-11-01 | 上海理工大学 | Retire test macro and method in a kind of Solid Rocket Engine Test three-dimensional combustion face |
| CN111751484A (en) * | 2020-04-13 | 2020-10-09 | 中国科学院力学研究所 | Solid-liquid rocket engine fuel burning rate measuring system |
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