Rocket flight result rapid evaluation method

Abstract

According to the rocket flight result rapid evaluation method, rocket flight data can be processed, an interpretation report is automatically generated, and working state rapid interpretation of pre-cooling before launching of a liquid rocket engine, fault diagnosis before ignition and in the flight process is achieved. According to the rapid judgment method for the rocket flight result, the analysis efficiency can be remarkably improved, and the time from the end of rocket flight to the obtaining of the engine analysis result is greatly shortened through automatic data processing, analysis process and rapid report generation. According to the evaluation method, the standard interpretation report is automatically generated, the problems of low manual analysis efficiency, incomplete analysis and the like are solved, and powerful support is provided for evaluation and improvement of rocket launching tasks.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of space launch vehicles, in particular to a rocket flight result rapid evaluation method. BACKGROUND

[0002] In a liquid rocket launch task, the engine as the core power source directly determines the success or failure of the task. During the rocket launch flight process, a large amount of data will be generated by each component of the engine, covering multiple dimensions of parameters such as pressure, temperature, and vibration frequency of key components. Post-launch system analysis is essential, and traditional analysis methods rely heavily on manual operation. Technical personnel need to screen, process and analyze massive data one by one, which not only consumes a lot of time and labor cost, but also is prone to data processing errors due to human negligence or fatigue, thereby causing deviations in the analysis results.

[0003] With the vigorous development of the aerospace industry, launch tasks are becoming more frequent. After each launch, it is urgently needed to obtain an accurate analysis report of the engine working results within a short time in order to quickly summarize lessons learned and optimize subsequent launch plans. However, the existing analysis mode is inefficient, and it often takes a long time from data collection to report generation, which makes it difficult to meet the demand for rapid confirmation of flight status.

[0004] Therefore, there is an urgent need to develop an efficient and accurate liquid rocket engine working condition rapid analysis method. SUMMARY

[0005] To solve the above technical problems, the present application provides a rocket flight result rapid evaluation method, which realizes the rapid interpretation of the working state of the liquid rocket engine before pre-cooling, before ignition, fault diagnosis and during flight, generates a standard interpretation report, solves the problems of low efficiency and incomplete analysis caused by human analysis, and provides strong support for the evaluation and improvement of rocket launch tasks.

[0006] The present application provides a rocket flight result rapid evaluation method, which processes the rocket flight data and automatically generates an interpretation report. The evaluation method of the present application at least includes the following steps:

[0007] Step one, obtain the relevant parameters of each engine pre-cooling process before rocket flight, and process them graphically to generate dynamic change curves;

[0008] Step two, obtain the relevant parameters of each engine ignition-1s to ignition time before rocket flight, process them into first average values respectively, and determine whether each first average value meets the ignition condition;

[0009] Step three, obtain the relevant parameters of the fault diagnosis time interval of each engine starting process, process them into second average values respectively, and determine whether each second average value meets the starting condition;

[0010] Step four, obtain the relevant parameters of each engine ignition after 4s to the steady state segment when the moment of shutdown, and perform graphical processing to generate dynamic change curve;

[0011] Step five, process the relevant parameters of the steady state segment into the third average value, and determine whether the third average value meets the stable working condition;

[0012] Step six, comprehensively determine the key parameters in the third average value to determine whether the engine is working normally;

[0013] Step seven, determine whether there is a step moment in the key parameters in the third average value, if there is, generate a dynamic change curve of the relevant parameters in the interval of ±2s of the step moment;

[0014] Step eight, according to the standardized report template, automatically generate a determination report for the above determination content.

[0015] Further, the method for comprehensively determining the key parameters in the third average value to determine whether the engine is working normally is: the thrust chamber oxygen injection pressure, the thrust chamber fuel injection pressure, the turbine inlet pressure and the turbine pump speed of each engine are determined one by one; if the number of abnormalities of the above four parameters in a certain submachine is ≥3, the information prompt of “the engine is working abnormally” is given in the report table, otherwise the information prompt of the number of abnormal parameters of the engine is given in the report table.

[0016] Further, the method for processing the relevant parameters of the steady state segment into the third average value and determining whether the third average value meets the stable working condition is: the oxygen inlet temperature, the oxygen inlet pressure, the oxygen pump pressure, the thrust chamber oxygen injection pressure, the fuel inlet temperature, the fuel inlet pressure, the fuel pump pressure, the thrust chamber fuel injection pressure, the turbine inlet pressure, the turbine inlet temperature, the turbine pump speed and the pump isolation cavity pressure of the steady state segment of the engine are processed into average values respectively, and are compared with the engine test theoretical parameter range at the time of delivery test in turn; if it is satisfied, the note “parameter normal” is given in the report chart, otherwise the note “parameter abnormal” is highlighted in the chart.

[0017] Further, after the relevant parameters of the steady state segment are processed into the third average value and it is determined whether the third average value meets the stable working condition, the engine thrust calculation is further included, specifically:

[0018] According to the formula: F=Y1×thrust chamber oxygen injection pressure+Pa×S, the thrust of each engine in the process of rocket flight is calculated and automatically filled into the report table; wherein Y1= test thrust / test thrust chamber fuel injection pressure, Pa is the environmental pressure at the flight altitude, and S is the nozzle exit area.

[0019] Furthermore, the engine thrust calculation further includes the following steps:

[0020] Step 1: Calculate the relationship between the liquid level and volume based on the liquid level at the start and end of the steady-state phase and the previous tank calibration characteristics;

[0021] Step 2: Calculate the propellant density in the steady-state section based on the average values ​​of the propellant temperature and pressure in the tank;

[0022] Step 3: Based on the relationship between the liquid level and the volume, the oxygen consumption flow rate and the fuel consumption flow rate are obtained using the volume change and the density;

[0023] Step 4: Get the engine mixture ratio and automatically fill in the report form.

[0024] Furthermore, the method of determining whether there is a step moment in the key parameters in the third average value, and if so, generating a dynamic change curve of the relevant parameters in the interval ±2s of the step moment, is specifically as follows:

[0025] Step 1: Use the cumulative sum algorithm to process the parameters of the turbine pump speed, turbine inlet pressure, thrust chamber oxygen injection pressure, and thrust chamber fuel injection pressure in the steady-state interval, with every N collected results as a unit;

[0026] Step 2: Use the formula: u i =u i-1 +(x i -x i-1 ) / N, calculate the current unit average value u i , through the formula: S i =|max(0, S i-1 +(x i -u i ))|, calculate the current unit cumulative sum S i ; where x i is the current value, x i-1 is the value at the previous moment, u i-1 is the average value of the previous unit, S i-1 Cumulative sum for the previous unit;

[0027] Step 3: Compare S i And given threshold H, if S i >H, it indicates that there is a step point at that moment, and the dynamic change curves of oxygen inlet temperature, oxygen inlet pressure, oxygen pump back pressure, thrust chamber oxygen injection pressure before, fuel inlet temperature, fuel inlet pressure, fuel pump back pressure, thrust chamber injection pressure before, turbine inlet pressure, turbine inlet temperature, turbine pump speed and pump isolation chamber pressure in the ±2s range of that moment are generated.

[0028] Further, the rocket flight result rapid evaluation method of the embodiment further comprises: acquiring relevant parameters in the interval of each engine ignition time to 4s, performing graphical processing thereon, generating a dynamic change curve, and generating a start transition state parameter satisfaction condition.

[0029] Further, the method for generating the start transition state parameter satisfaction condition is: processing the average value X1 of the oxygen pre-injection pressure in the thrust chamber from 3.8s to 4s after each engine sub-engine is ignited, finding the time point at which the oxygen pre-injection parameter in the start process first reaches the X1*90% value, and recording the time point in the transition state performance table.

[0030] In one embodiment, the rocket flight result rapid evaluation method of the embodiment further comprises: acquiring the oxygen pre-injection pressure parameter in the thrust chamber from-2s before the engine is shut down to the shutdown time, generating a dynamic change curve, and processing the oxygen pre-injection pressure parameter into an average value X2, finding the time point at which the oxygen pre-injection parameter in the shutdown process first reaches the X2*5% value, and recording the time point as a shutdown deceleration parameter.

[0031] In any one of the above embodiments, the relevant parameters in the acquisition of the relevant parameters in the pre-cooling process of each engine before the rocket flight at least include: oxygen inlet temperature, oxygen inlet pressure, oxygen pre-cooling backflow valve front temperature, fuel inlet temperature, fuel inlet pressure, fuel pre-cooling backflow valve front temperature, control gas cylinder pressure, purge gas cylinder pressure, and pump isolation cavity pressure.

[0032] The rocket flight result rapid evaluation method provided by the application realizes rapid interpretation of the working state of the liquid rocket engine before launch, before ignition, fault diagnosis and in the flight process, and generates a standard interpretation report, solves the problems of low analysis efficiency and incomplete analysis, and provides strong support for evaluation and improvement of rocket launch tasks.

[0033] The rocket flight result rapid evaluation method of the application can intelligently obtain an interpretation report without human influence. While ensuring the accuracy and precision of the extracted parameters, the analysis result has higher objectivity and comparability. In addition, this evaluation method is based on the original data processing and calculation of rocket flight, can automatically generate a format-standard interpretation report, can visualize various contents of the test report, and can effectively display the horizontal comparison result.

[0034] Those skilled in the art will recognize additional features and advantages upon reading the detailed description and viewing the drawings. BRIEF DESCRIPTION OF DRAWINGS

[0035] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed in the embodiments. Obviously, the drawings in the following description only constitute some embodiments of the present application, and for those skilled in the art, other drawings can also be obtained from these drawings without creative effort.

[0036] Figure 1 is a flow chart of the operation process of the rocket flight result rapid evaluation method of the embodiment of the present application.

[0037] Figure 2 is a specific operation flow chart of the rocket flight result rapid evaluation method of the embodiment of the present application.

[0038] Figure 3 is an operation flow chart of the embodiment of the present application for determining whether there is a step moment in the key parameters in the third average value. DETAILED DESCRIPTION

[0039] The features and exemplary embodiments of various aspects of the present application will be described in detail below, in order to make the purposes, technical solutions and advantages of the present application more clear and apparent, the following will further describe the present application in combination with the drawings and specific embodiments. It should be understood that the specific embodiments described herein are only configured to explain the present application, for exemplary illustration of the principles of the present application, and are not configured to limit the present application. In addition, the structural members in the drawings are not necessarily drawn to scale. For example, the size of some structural members in the drawings can be enlarged for other structural members or regions, to help understand the embodiments of the present application.

[0040] The orientation words appearing in the following description are the directions shown in the drawings, and are not limited to the specific structure of the embodiments of the present application. In the description of the present application, it should be noted that, unless otherwise specified, the terms "mounting", "connecting", "connecting" should be understood broadly, for example, it can be fixedly connected, or it can be detachably connected, or integrally connected; it can be directly connected, or indirectly connected through an intermediate medium. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0041] In addition, the terms "including", "containing", "having" or any other variants thereof are intended to cover non-exclusive inclusion, so that the inclusion of a series of element structures or components not only includes those elements, but also includes other structures or components which are not explicitly listed or inherent in the structures or components. Without more limitation, the elements defined by the statement "including" do not exclude the presence of other same elements in the article or device including the elements.

[0042] Spatially relative terms such as "under", "below", "lower", "above", "upper", and the like are used for ease of description to explain the positioning of one element relative to a second element. With respect to the devices illustrated, the terms are intended to encompass different orientations of the device and different orientations of the components of the device. For example, if a device described includes upper and lower components, the device can also be implemented with the upper and lower components in different positions, e.g., the upper component can be below the lower component. Such terms as "first" and "second" are also used to describe various elements, regions, sections, etc. and are not always used literally. Like terms are used to describe like elements, regions, sections, etc.

[0043] The present application can be implemented without some of these specific details. The following description of the embodiments is merely exemplary in nature and is provided to give a better understanding of the present application.

[0044] Referring to Figure 1 The rocket flight result rapid evaluation method provided by the embodiments of the present application can process rocket flight data to automatically generate an interpretation report, and at least includes: generating a pre-cooling process engine related temperature and pressure curve, generating a pre-ignition engine minimum launch condition parameter satisfaction, generating a start-up process fault diagnosis parameter satisfaction, generating an engine parameter steady-state segment working curve, processing and judging each parameter in the steady-state segment, multi-parameter comprehensive judgment, engine performance parameter processing, flight process abnormal situation judgment and recording, generating a start-up segment parameter working curve, generating a start-up transition state parameter satisfaction, generating a shutdown segment parameter working curve, generating a shutdown transition state parameter satisfaction, and the like, and at least automatically generating a judgment report for the above corresponding parameter curves and judgment results.

[0045] It should be noted that engine ignition and engine start-up are different states, and engine ignition only refers to the moment of ignition, while engine start-up refers to the stage after engine ignition is completed.

[0046] Referring to Figure 2 Specifically, the rocket flight result rapid evaluation method provided by the present application at least includes the following steps:

[0047] S100, acquiring related parameters in a pre-cooling process of each engine before rocket flight, and performing graphical processing to generate a dynamic change curve;

[0048] S200, acquiring related parameters of each engine from ignition-1s to ignition moment before rocket flight, processing the related parameters into first average values respectively, and judging whether each first average value satisfies an ignition condition;

[0049] S300, acquire the related parameters of each engine starting process fault diagnosis time interval, process them into second average values respectively, and determine whether each second average value meets the starting condition;

[0050] S400, acquire the related parameters of the steady state section from 4s after each engine ignition to the shutdown moment, and process them graphically to generate dynamic change curves;

[0051] S500, process the related parameters of the steady state section into third average values, and determine whether the third average values meet the stable working condition;

[0052] S600, comprehensively determine the key parameters in the third average values to determine whether the engine works normally;

[0053] S700, determine whether there is a step moment in the key parameters in the third average values, if there is, generate a dynamic change curve of the related parameters in the ±2s interval of the step moment;

[0054] S800, automatically generate a determination report according to a standardized report template based on the above determination contents.

[0055] Further, S100, the method for acquiring the related parameters in the pre-cooling process of each engine before the rocket flight and processing them graphically to generate dynamic change curves is: according to the rocket flight data, acquiring the parameters such as oxygen inlet temperature, oxygen inlet pressure, oxygen pre-cooling backflow valve temperature, fuel inlet temperature, fuel inlet pressure, fuel pre-cooling backflow valve temperature, control gas cylinder, blowing gas cylinder pressure and pump isolation cavity pressure in the pre-cooling process of each engine, and processing them graphically respectively to generate dynamic change curves of each parameter, and further acquiring the parameter change and characteristics in the pre-cooling process.

[0056] Further, S200, the method for acquiring the related parameters from ignition-1s to ignition moment of each engine before the rocket flight and processing them into first average values, and determining whether each first average value meets the ignition condition is: according to the rocket flight data, acquiring the parameters such as oxygen inlet temperature, oxygen inlet pressure, oxygen pre-cooling backflow valve temperature, fuel inlet temperature, fuel inlet pressure, fuel pre-cooling backflow valve temperature, control gas cylinder pressure, blowing gas cylinder pressure, pump isolation cavity pressure of each engine from ignition-1s to ignition moment, and processing them into first average values respectively. Compare the first average values of each parameter with the set minimum launch condition in turn, if it meets, add a note "meets" in the report chart, otherwise, add a note "does not meet" in the report chart.

[0057] Further, the method for obtaining the related parameters of each engine starting process fault diagnosis time interval, processing them into second average values respectively, and determining whether each second average value meets the starting condition is as follows: according to the rocket flight data, the thrust chamber oxygen injection pre-pressure, the thrust chamber fuel injection pre-pressure, the turbine pump inlet pressure and the turbine pump speed parameters of the engine starting process fault diagnosis time interval are obtained and processed into average values respectively. Each average value is compared with the set fault diagnosis parameter range interval in turn. If it meets the condition, “meets” is noted in the chart. Otherwise, “does not meet” is noted in the chart in bold and red.

[0058] Further, the method for obtaining the related parameters of each engine starting process fault diagnosis time interval, processing them into second average values respectively, and determining whether each second average value meets the starting condition is as follows: according to the rocket flight data, the thrust chamber oxygen injection pre-pressure, the thrust chamber fuel injection pre-pressure, the turbine pump inlet pressure and the turbine pump speed parameters of the engine starting process fault diagnosis time interval are obtained and processed into average values respectively. Each average value is compared with the set fault diagnosis parameter range interval in turn. If it meets the condition, “meets” is noted in the chart. Otherwise, “does not meet” is noted in the chart in bold and red.

[0059] Further, the method for obtaining the related parameters of each engine starting process fault diagnosis time interval, processing them into second average values respectively, and determining whether each second average value meets the starting condition is as follows: according to the rocket flight data, the thrust chamber oxygen injection pre-pressure, the thrust chamber fuel injection pre-pressure, the turbine pump inlet pressure and the turbine pump speed parameters of the engine starting process fault diagnosis time interval are obtained and processed into average values respectively. Each average value is compared with the set fault diagnosis parameter range interval in turn. If it meets the condition, “meets” is noted in the chart. Otherwise, “does not meet” is noted in the chart in bold and red.

[0060] Further, the method for obtaining the related parameters of each engine starting process fault diagnosis time interval, processing them into second average values respectively, and determining whether each second average value meets the starting condition is as follows: according to the rocket flight data, the thrust chamber oxygen injection pre-pressure, the thrust chamber fuel injection pre-pressure, the turbine pump inlet pressure and the turbine pump speed parameters of the engine starting process fault diagnosis time interval are obtained and processed into average values respectively. Each average value is compared with the set fault diagnosis parameter range interval in turn. If it meets the condition, “meets” is noted in the chart. Otherwise, “does not meet” is noted in the chart in bold and red.

[0061] Further, the related parameters of the steady state section are processed as a third average value, and after determining whether the third average value meets the stable working condition, engine thrust calculation is further included, specifically: the steady state section parameters (third average value) processed by S500 are further processed, the thrust coefficient Y1 is obtained according to the ground acceptance test, Y1 = test thrust / test thrust chamber fuel pre-injection pressure; then the thrust of each engine in the rocket flight process is calculated according to the formula: F = Y1 × thrust chamber oxygen pre-injection pressure third average value + Pa × S, and is automatically filled into the report form, and then the thrust characteristics in the whole flight process are obtained; wherein Y1 is the thrust coefficient obtained during the ground acceptance test, Pa is the environmental pressure at the flight height, S is the nozzle exit area, and F is the thrust in the rocket flight process.

[0062] Further, after the engine thrust calculation, the following steps are further included: according to the liquid level height at the start and end time of the steady state section and the previous tank calibration characteristics, the relationship between the liquid level height and the volume is calculated; according to the average value of the tank propellant temperature and pressure in the steady state section, the propellant density in the section is calculated; according to the relationship between the liquid level height and the volume, the oxygen consumption flow and the fuel consumption flow are obtained by using the volume change and the density respectively; the engine mixture ratio is obtained, and is automatically filled into the report form.

[0063] Referring to Figure 3 , further, S700, determining whether there is a step time in the third average value, if there is, the method for generating the dynamic change curve of the related parameters in the step time ± 2s interval is:

[0064] S701, using the cumulative sum algorithm to process the turbine pump speed, the turbine inlet pressure, the thrust chamber oxygen pre-injection pressure and the thrust chamber fuel pre-injection pressure parameters in the steady state section interval, taking every N acquisition results as a unit;

[0065] S702, calculating the current unit average value u i by the formula: u i-1 = u i + (x i-1 -x i ) / N, and calculating the current unit cumulative sum S i by the formula: S i-1 = |max (0, S i + (x i -u i )) |; wherein x i is the current time value, x i-1 is the previous time value, u i-1 is the previous unit average value, and S i-1 is the previous unit cumulative sum;

[0066] S703, comparing S iand a given threshold H, if S i >H, it indicates that there is a step point at this moment, and the dynamic change curves of oxygen inlet temperature, oxygen inlet pressure, oxygen pump post-pressure, thrust chamber oxygen injection pre-pressure, fuel inlet temperature, fuel inlet pressure, fuel pump post-pressure, thrust chamber fuel injection pre-pressure, turbine inlet pressure, turbine inlet temperature, turbine pump speed and pump isolation cavity pressure in the interval of this moment ±2s are generated.

[0067] The specific steps are as follows: the cumulative sum algorithm is used to process the turbine pump speed, turbine inlet pressure, thrust chamber oxygen injection pre-pressure and thrust chamber fuel injection pre-pressure data in the steady state interval, and every N acquisition results are taken as a unit, the current time value is denoted as x i , the previous time value is denoted as x i-1 , u i- 1 is the average value of the previous unit, u i is the average value of the current unit, wherein u i =u i-1 +(x i -x i-1 ) / N. S i-1 is the cumulative sum of the previous unit, S i is the cumulative sum of the current unit, wherein S i =|max(0, S i-1 +(x i -u i ))|. S i is compared with a given threshold H, if S i >H (given threshold), it indicates that there is a step point at this moment, and the parameter curves of each parameter (oxygen inlet temperature, oxygen inlet pressure, oxygen pump post-pressure, thrust chamber oxygen injection pre-pressure, fuel inlet temperature, fuel inlet pressure, fuel pump post-pressure, thrust chamber fuel injection pre-pressure, turbine inlet pressure, turbine inlet temperature, turbine pump speed and pump isolation cavity pressure) of the engine in the interval of this moment ±2s are automatically drawn to provide a reference for subsequent detailed analysis.

[0068] Further, the rocket flight result rapid evaluation method of the embodiment of the present application further comprises generating a starting segment parameter working curve and a starting transition state parameter satisfaction condition, and the specific steps are as follows: according to the rocket flight data, the related parameters (engine oxygen inlet temperature, oxygen inlet pressure, oxygen pump post-pressure, thrust chamber oxygen injection pre-pressure, fuel inlet temperature, fuel inlet pressure, fuel pump post-pressure, thrust chamber fuel injection pre-pressure, turbine inlet pressure, turbine inlet temperature, turbine pump speed, pump isolation cavity pressure) of each sub-machine in the interval from the ignition moment to 4s are obtained, and dynamic change curves are respectively generated. The thrust chamber oxygen injection pre-pressure of each sub-machine from 3.8s to 4s after ignition is processed as an average value X1, and the time point at which the thrust chamber oxygen injection pre-pressure reaches X1*90% for the first time in the starting process is found, and the time point is recorded in the transition state performance table.

[0069] Further, the rocket flight result rapid evaluation method of the embodiment of the present application further comprises generating a shutdown section parameter working curve and generating a shutdown transition state parameter satisfaction condition, and the specific steps are as follows:

[0070] According to the rocket flight data, each parameter (oxygen inlet pressure, oxygen pump rear pressure, thrust chamber oxygen injection front pressure, fuel inlet pressure, fuel pump rear pressure, thrust chamber fuel injection front pressure, turbine inlet pressure, turbine inlet temperature, turbine pump speed, pump isolation cavity pressure) of each sub-machine in the interval from the shutdown moment to 6s after the shutdown is obtained, and a dynamic change curve is generated respectively.

[0071] The thrust chamber oxygen injection front pressure parameter from-2s before the engine shutdown to the shutdown moment is obtained, and a dynamic change curve is generated.

[0072] The thrust chamber oxygen injection front pressure parameter from-2s before the engine shutdown to the shutdown moment is processed as an average value X2, the moment when the thrust chamber oxygen injection front parameter first reaches the value of X2*5% in the shutdown process is found, and the moment is recorded in the transition state performance table and used as the shutdown deceleration parameter.

[0073] The abnormal result in any one of the above embodiments is stored in a fault database, and each time of launch data can be compared with historical fault data. For example, the fault occurring again will be highlighted and the historical fault data will be displayed as a reference.

[0074] Finally, according to a standardized report template, the parameter curves and evaluation results are automatically generated to generate a determination report.

[0075] The above embodiments can be combined with each other and have corresponding technical effects.

[0076] The rocket flight result rapid evaluation method of the present application can significantly improve the analysis efficiency. Through the automatic data processing, analysis process and rapid report generation, the time from the end of the rocket flight to the analysis result of the engine is greatly shortened. Compared with the traditional manual analysis method, the analysis efficiency is improved by several times or even dozens of times, and a high-quality analysis report can be quickly generated within a few minutes after the rocket launch, which provides support for the rapid evaluation of the flight result.

[0077] The rocket flight result rapid evaluation method of the present application can significantly improve the data analysis integrity. Through the graphical visualization display, steady-state data and step point processing, the parameter changes in the whole flight process are covered, and the engine working state is more completely characterized.

[0078] The above only describes the preferred embodiments of the present application and does not limit the present application. Any modification, equivalent replacement, improvement, etc. made within the spirit and principles of the present application shall be included in the protection scope of the present application.

Claims

1. A method for rapid evaluation of rocket flight results, which processes rocket flight data and automatically generates an interpretation report, characterized in that: At least the following steps are included: Step 1: Obtain relevant parameters of each engine during the pre-cooling process before the rocket flight, and perform graphical processing on them to generate a dynamic change curve; Step 2: Obtain relevant parameters of each engine from ignition -1s before the rocket flight to the ignition time, process them into first average values, and determine whether each first average value meets the ignition condition; Step 3: Obtain relevant parameters of the fault diagnosis time interval of each engine starting process, process them into second average values, and determine whether each second average value meets the starting condition; Step 4: Obtain relevant parameters of the steady-state section from 4 seconds after ignition to the shutdown moment of each engine, and perform graphical processing on them to generate a dynamic change curve; Step 5: Process the relevant parameters of the steady-state section into a third average value, and determine whether the third average value meets the stable working conditions; Step 6: Comprehensively determine the key parameters in the third average value to determine whether the engine is operating normally; Step 7: Determine whether there is a step moment in the key parameters in the third average value. If so, generate a dynamic change curve of the relevant parameters in the interval of ±2s of the step moment; Step 8: Automatically generate a judgment report based on the above judgment contents according to the standardized report template.

2. The rocket flight result rapid evaluation method according to claim 1, characterized in that: The method for comprehensively judging the key parameters in the third average value to determine whether the engine is working normally is as follows: Comprehensively assess each engine's thrust chamber oxygen pre-injection pressure, thrust chamber fuel pre-injection pressure, turbine inlet pressure, and turbine pump speed. If the number of abnormalities in the above four parameters in a certain extension is ≥3, the information prompt "the engine is working abnormally" will be given in the report form; otherwise, the information prompt of the number of abnormal parameters of the engine will be given in the report form.

3. The rocket flight result rapid evaluation method according to claim 2, characterized in that: The method of processing the relevant parameters of the steady-state section into a third average value and determining whether the third average value meets the stable working condition is: The steady-state engine oxygen inlet temperature, oxygen inlet pressure, oxygen pump outlet pressure, thrust chamber oxygen injection pressure, fuel inlet temperature, fuel inlet pressure, fuel pump outlet pressure, thrust chamber fuel injection pressure, turbine inlet pressure, turbine inlet temperature, turbine pump speed, and pump isolation chamber pressure are processed as average values ​​and compared with the theoretical parameter range of the engine test at the time of delivery test. If the conditions are met, then note "parameters normal" in the report chart; otherwise, highlight "parameters abnormal" in the chart.

4. The method for rapid evaluation of rocket flight results according to claim 3, characterized in that: The processing of the relevant parameters of the steady-state section into a third average value and determining whether the third average value meets the stable working condition also includes engine thrust calculation, specifically: According to the formula: F = Y1 × the third average value of the thrust chamber oxygen pressure before injection + Pa × S, the thrust of each engine during the rocket flight is calculated and automatically filled in the report form; Where, Y1 = test thrust / pressure in the test thrust chamber before combustion, Pa is the ambient pressure at the flight altitude, and S is the nozzle outlet area.

5. The rocket flight result rapid evaluation method according to claim 4, characterized in that: The engine thrust calculation also includes: Calculate the relationship between liquid level and volume based on the liquid level at the start and end of the steady-state phase and the previous tank calibration characteristics; Calculate the propellant density in the steady-state section based on the average value of the tank propellant temperature and pressure in the steady-state section. According to the relationship between liquid level and volume, the oxygen consumption flow rate and fuel consumption flow rate are obtained respectively using volume change and density; Get the engine mixture ratio and automatically fill in the report form.

6. The method for rapidly evaluating rocket flight results according to claim 1, wherein the method for determining whether a key parameter in the third average value has a step moment and, if so, generating a dynamic change curve of the relevant parameter within the interval ±2s of the step moment is specifically: The cumulative sum algorithm is used to process the parameters of the turbine pump speed, turbine inlet pressure, thrust chamber oxygen injection pressure, and thrust chamber fuel injection pressure in the steady-state interval, with every N collected results as a unit. By formula: i =u i-1 +(x i -x i-1 ) / N, calculate the current unit average value u i , through the formula: S i =|max(0, S i-1 +(x i -u i ))|, calculate the current unit cumulative sum S i ;in, x i is the current value, x i-1 is the value at the previous moment, u i-1 is the average value of the previous unit, S i-1 Cumulative sum for the previous unit; Compare S i And given threshold H, if S i >H, it indicates that there is a step point at that moment, and the dynamic change curves of oxygen inlet temperature, oxygen inlet pressure, oxygen pump back pressure, thrust chamber oxygen injection pressure before, fuel inlet temperature, fuel inlet pressure, fuel pump back pressure, thrust chamber injection pressure before, turbine inlet pressure, turbine inlet temperature, turbine pump speed and pump isolation chamber pressure in the ±2s range of that moment are generated.

7. The method for rapid evaluation of rocket flight results according to claim 1, characterized in that: Also includes: The relevant parameters of each engine from the ignition moment to the 4s interval are obtained, and graphically processed to generate a dynamic change curve and the starting transition state parameter satisfaction status.

8. The method for rapid evaluation of rocket flight results according to claim 7, characterized in that: The method for generating the starting transition state parameter satisfaction condition is: Process the average thrust chamber oxygen injection pressure X1 from 3.8s to 4s after ignition of each engine branch, find the time point when the thrust chamber oxygen injection parameter first reaches the value of X1*90% during the startup process, and record this time point in the transient state performance table.

9. The method for rapid evaluation of rocket flight results according to claim 1, characterized in that: Also includes: Obtain the thrust chamber oxygen pre-injection pressure parameters from -2 seconds before engine shutdown to the moment of shutdown, and generate a dynamic change curve; The thrust chamber oxygen pre-injection pressure parameter is processed into the average value X2, and the moment when the thrust chamber oxygen pre-injection parameter first reaches the value of X2×5% during the shutdown process is found. This moment is recorded and used as the shutdown deceleration parameter.

10. The method for rapid evaluation of rocket flight results according to claim 1, characterized in that: The relevant parameters obtained during the pre-cooling process of each engine before the rocket flight include at least: oxygen inlet temperature, oxygen inlet pressure, temperature before the oxygen pre-cooling return valve, fuel inlet temperature, fuel inlet pressure, temperature before the fuel pre-cooling return valve, control gas cylinder pressure, blow-off gas cylinder pressure and pump isolation chamber pressure.

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