Aircraft engine combustion chamber casing burn-through test method, system and device

Abstract

The invention relates to the technical field of ground testing of aircraft engines, and discloses a burn-through test method, system and device for an aircraft engine combustion chamber casing, which simulates the combustion environment in which flames are ejected from the aircraft engine combustion chamber casing and ablates protective parts in a flight environment; a plurality of data acquisition tools and protective parts are switched in a detachable installation manner to obtain flame characteristic data and protective part ablation data respectively; at least one of the data acquisition tools obtains the three-dimensional distribution of corresponding flame characteristic data in a circumferential multi-layer and multi-point data acquisition manner; the scheme studies the influence of three-dimensional flames on different types of protective parts, which is conducive to analyzing the relationship between the data, especially studying the relationship between the burn-through degree of the protective parts and the combustion conditions of the engine, so as to obtain a more complete judgment on the performance of the protective parts, and promote the precise installation position of new more three-dimensional protective parts.

Description

Technical Field

[0001] The invention relates to the technical field of ground testing of aircraft engines, and in particular to a method, system and device for burning-through testing of a combustion chamber casing of an aircraft engine. Background Art

[0002] The combustion chamber casing of an aircraft engine may burn through from the inside, causing the casing to rupture. The high-temperature and high-pressure gas inside the combustion chamber will be ejected from the rupture, forming a high-speed, high-temperature and high-pressure jet fire, causing irreversible damage to the engine. In order to protect the important external components of the engine combustion chamber, protective parts are placed on the outside of the location where the casing may burn through. Therefore, in order to evaluate the performance of engine protective parts in extreme thermal environments, the combustion chamber casing burn-through test of aircraft engines is one of the important tests, and the industry is also very concerned about the research progress of this burn-through test.

[0003] At present, there is no standard test method for aircraft engine burn-through test in China. The conventional test method is to determine the flame parameters according to relevant foreign test research and Federal Aviation Administration (FAA) regulations to simulate the jet fire at the moment of casing burn-through, and then place the protective baffle at the specified distance, and judge and select the performance of the protective baffle based on whether it burns through. This method was feasible in the past when the flight requirements were not high, the flight environment was simple, and the aviation technology was not mature enough. However, with the continuous improvement of aviation flight safety requirements, the increasingly changeable flight environment, and the continuous innovation of aviation technology, if the previous conventional method is still used to conduct the performance test of protective parts in the existing more complex flight environment, there are limitations. Although the error can be minimized as much as possible through repeated tests, it will still affect the effective and accurate evaluation of the real performance of new protective parts in complex flight environments.

[0004] In order to meet these challenges, it is an inevitable trend to optimize the test methods simultaneously. For example, the Chinese document "Study on the Burn-through Flame Characteristics of the Combustion Chamber of an Aircraft Engine" studied the influence of nozzle aperture, oil-gas mass ratio, outer air velocity, and internal pressure of the combustion chamber on the characteristics of the jet fire, but it only studied the axial change characteristics of the flame ejection path. For complex flight environments, this method cannot analyze the true characteristics of the flame in the complex flight environment of high altitude, high pressure and high temperature, and therefore cannot accurately evaluate the performance of the corresponding protective parts. In response to the problem of the test method of single axial motion, Chinese patent CN115453039A discloses a supersonic flame ablation test device and method for ablation-resistant materials, which proposes to change the flame ejection path to form a dynamic, variable angle flame to meet the requirements of multi-angle and multi-attitude ablation tests, and provides a number of fixed parameter data, but it only changes the path of the flame ejection, and the entire test idea has not changed in essence. The configuration parameters of the flame, the placement of the protective plate at a specified distance, and the judgment and selection of the performance of the protective plate based on whether it is burned through are the same as the traditional method, and there is no more data value for the improvement of how the protective parts cope with complex flight environments. Summary of the invention

[0005] The present invention aims to provide a burn-through test method, system and device for an aircraft engine combustion chamber casing. Through the test of this scheme, the performance of different protective parts in a complex flight environment can be evaluated more quickly, accurately and truly, which is conducive to studying the optimization direction of different types of protective parts in response to existing challenges.

[0006] The basic scheme provided by the present invention is: a burn-through test method for a combustion chamber casing of an aircraft engine, the method comprising:

[0007] Step S100, simulating a combustion environment in which flames are ejected from a combustion chamber casing of an aircraft engine and ablates protective components in a flight environment;

[0008] Step S200, flame characteristic test:

[0009] S201, within a preset time after the flame is ejected with the determined configuration parameters, each flame data acquisition tool starts to move away from the flame ejection starting point in sequence according to the axial position change speed, and performs flame data acquisition during the axial position change process; at least one of the data acquisition tools acquires the corresponding flame characteristic data at each axial position in a circumferential multi-layer and multi-point data acquisition manner, and aggregates to obtain a three-dimensional distribution of the corresponding flame characteristic data;

[0010] S202, each flame data acquisition tool is switched in a detachable installation manner, and the data acquired by all flame data acquisition tools are aggregated to obtain the flame characteristic data under the configuration parameters and axial position change speed;

[0011] S203, adjusting the configuration parameters and the axial position change speed, and repeating the above steps to obtain corresponding flame characteristic data;

[0012] Step S300, performing a burn-through test of the protective component:

[0013] S301, determining the axial distance between the flat test fixture and the flame ejection starting point based on the flame characteristic data obtained from the configuration parameters and the axial position change speed in S200, and obtaining the historical experience axial distance to obtain an axial distance set;

[0014] S302, within a preset time after the flame is sprayed with the configuration parameters in S200, the non-curved protective parts are installed in sequence using the flat test fixture, and a burn-through test is performed at all axial distances in the axial distance set to obtain the protective part ablation data corresponding to the configuration parameters and the axial distance set;

[0015] S303, based on the shield ablation data obtained from the shield without curvature in S302, determine the axial offset angle between the flat test fixture and the flame ejection starting point and the basic parameter curvature radius of the shield, adjust the axial offset angle and the curvature radius of the shield, and repeat S302 to obtain the corresponding shield ablation data;

[0016] S304, adjusting other basic parameters of the protective element, such as material thickness and material type, and repeating the above steps to obtain corresponding protective element ablation data;

[0017] Step S400: Based on the flame characteristic data and the shield ablation data, the relationship between the flame configuration parameters, the shield basic parameters, the axial distance, and the axial offset angle is analyzed.

[0018] The invention is based on a burn-through test method for a combustion chamber casing of an aero-engine and also provides a burn-through test device for a combustion chamber casing of an aero-engine.

[0019] The device comprises a base and a flame temperature measuring tool, a flame heat flux density measuring tool, a flame impact force measuring tool and a flat plate test tool which are detachably mounted on the base;

[0020] The flame temperature measuring tool comprises a first mounting plate and a plurality of sensors; the first mounting plate is provided with a first central hole for the flame to pass through; a plurality of sensor mounting holes are radially provided at preset positions along the outer circumference of the central hole, wherein the number and positions of the sensors installed in all the sensor mounting holes and the relative distances of the sensors into the sensor mounting holes correspond to the circumferential multi-layer and multi-point digital sampling method;

[0021] The flame heat flux density measuring tool and the flame impact force measuring tool both include a second mounting plate; the second mounting plate is provided with a second center hole for the flame to pass through; at least one heat flux meter is installed in the second center hole of the flame heat flux density measuring tool; at least one pressure sensor and its isolation component are installed in the second center hole of the flame impact force measuring tool; the flat plate test tool includes a third mounting plate for installing a protective member;

[0022] Wherein, the first mounting plate, the second mounting plate and the third mounting plate are all mounted with a first special-shaped structure and a second special-shaped structure for detachable installation.

[0023] The present invention also provides an aircraft engine combustion chamber casing burn-through test system, including a gas generator and a nitrogen supply system, an air supply system, an oxygen system, a kerosene supply system, a water supply system, an automatic gas distribution system and a measurement and control system connected to the gas generator, which is used to simulate the combustion environment of the aircraft engine combustion chamber casing in a flight environment where flames are ejected and protective parts are burned, and to form different casing burn-through flames and their corresponding configuration parameters; it also includes an aircraft engine combustion chamber casing burn-through test device, which is placed at the front end of the gas generator flame ejection, and is used to collect flame characteristic data and protective part ablation data.

[0024] The working principle and advantages of the present invention are as follows: in the existing burn-through test of the combustion chamber casing of an aircraft engine, the conventional test judges the performance by whether the protective plate is burned through. Based on the result of burn-through and the high temperature, high speed and high pressure characteristics of the ejected flame, the conventional method is to take the flame center with the largest energy as the research object. The test results study the ablation of a small area of ​​the protective plate. In view of the fact that the study is about the ablation range of a small area, the conventional test regards the protective plate as a plate-like structure. Therefore, in theory, it is feasible to conduct a burn-through judgment study based on the straight-line distance from the flame center to the protective plate. However, as the flight environment becomes more and more complex and aviation technology continues to develop, the flame will not be ejected in an absolutely predetermined direction in a complex flight environment with high temperature, high pressure or frequent switching of turbulent steady state. The ablation area and ablation effect are not limited to a small area of ​​ablation evenly spread from the flame center. The form of the protective part is not limited to the plate-like structure, and there may be more forms of new protective parts. At the same time, due to the high severity of burn-through in the combustion chamber of an aircraft engine, the data obtained from a single test and the performance of protective parts judged solely by whether they have been burnt through are far from meeting the existing safety protection research, especially the relationship between the flame characteristics in complex flight environments and the performance of protective parts in more forms, which is a key link in improving protection safety. In addition, there is currently no corresponding equipment in the industry to meet the needs of rapid and accurate acquisition of multi-dimensional data in burn-through tests.

[0025] In the face of existing test challenges, this program breaks through the conventional limited understanding of flames and protective parts in flight environments, fully recognizes the changes in the circumferential temperature of flames in flight environments, takes into account the three-dimensional characteristics of flames, the changes in the shape of protective parts, and the diversity of the causes and forms of combustion of aircraft engines. The study is about the impact of three-dimensional flames on different types of protective parts. In one test, the three-dimensional distribution data of flames can be obtained, which is not limited to judging the performance of protective parts based on whether they are burned through, but to obtain more dimensional protective parts and flame data. At the same time, for the test protective parts, frontal impact is the most severe state, but in actual use, the path of the frontal impact flame can be changed to reduce the frontal impact of the test piece and ensure that the test protective parts are not penetrated. Therefore, it is of practical significance to study inclined or curvature test protective parts. Through the test method of this scheme, on the basis of obtaining three-dimensional flame characteristic data, not only the ablation data of conventional non-curved protective parts are obtained, but also the changing relationship between the tilt offset angle and the curvature radius of the protective parts is constructed. Since the axial offset angle and the curvature radius of the protective parts used in the test are obtained from the current test data, the study of the data relationship is more valuable, and the specific parameter range that different material test pieces can withstand flames can be given. For example, a certain material can withstand the flame head-on at a certain distance from the starting point of the flame, or can withstand the flame within a certain range of tilt angles, or can withstand the flame when the curvature is greater than a certain value. In this way, a more complete and multi-angle performance judgment of the protective parts can be obtained, which is especially helpful for clarifying the precise installation position of new more three-dimensional protective parts, and studying the relationship between the degree of burn-through of the protective parts and the combustion conditions of the engine.

[0026] Compared with the prior art, based on the test method optimized in this scheme, the optimized device structure of this scheme ensures that the number and position arrangement of points can reflect the characteristics of three-dimensional flames in the flight environment, and it is easy to disassemble and assemble, and the positioning is fast. Compared with the test of simple axial distance, this scheme can obtain more data in the same test time, with a long service life and cost savings. Combined with the use of the system of this scheme, there are many adjustable parameters, which can simulate various engine combustion conditions, which is conducive to establishing the relationship between the ablation of protective parts and the combustion of the engine, and can realize the reverse judgment of the cause of engine combustion from the ablation degree of protective parts. In addition, the use of the system makes the test repeatable, which can ensure the consistency of the test conditions and effectively support the subsequent analysis of the relationship between test data. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 A flow chart of a burn-through test method for a combustion chamber casing of an aircraft engine provided by an embodiment of the present invention;

[0028] Figure 2 A schematic structural diagram of an aircraft engine combustion chamber casing burn-through test system provided by an embodiment of the present invention;

[0029] Figure 3 A schematic diagram of the structure of a base provided by an embodiment of the present invention;

[0030] Figure 4 A schematic diagram of the structure and use of a flame temperature measurement tool provided in an embodiment of the present invention;

[0031] Figure 5 A schematic diagram of the structure of a flame heat flux density measurement tool (flame impact force measurement tool) provided in an embodiment of the present invention;

[0032] Figure 6 A schematic diagram of the installation of a pressure sensor in a flame impact force measurement tool provided in an embodiment of the present invention;

[0033] Figure 7 A schematic diagram of the structure and use of a flat plate test tool provided in an embodiment of the present invention;

[0034] Figure 8 A schematic diagram of axial position change provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0035] The following is a further detailed description through specific implementation methods:

[0036] The marks in the drawings of the specification include: gas generator 1, torch igniter 11, base 2, cross bar 21, bottom plate 22, inclined support rod 23, flame temperature measuring tool 3, temperature sensor 31, first mounting plate 32, first center hole 33, sensor mounting hole 34, nut 35, flame impact force measuring tool 4, pressure sensor 41, second mounting plate 42, second center hole 43, ceramic sheet 44, composite plate 45, flat plate test tool 5, protective member 51, third mounting plate 52, curved protective member 53, external load 54, first special-shaped structure 6, upper arc 61, lower arc 62, second special-shaped structure 7.

[0037] Embodiment 1

[0038] Basically as attached Figure 1 , Figure 2 Shown: aircraft engine combustion chamber casing burn-through test device and system, used to execute aircraft engine combustion chamber casing burn-through test method.

[0039] like Figure 3 , Figure 4 , Figure 5 , Figure 6 and Figure 7As shown, the aircraft engine combustion chamber casing burn-through test device includes a base 2 and a flame temperature measurement tool 3, a flame heat flux density measurement tool, a flame impact force measurement tool 4 and a flat plate test tool 5 which are detachably mounted on the base 2.

[0040] In the actual burn-through situation, the flame ejection is affected by the space between the combustion chamber and the protective part. This space is not only determined by the straight-line distance, but also by the combustion chamber shell and the shape of the protective part. The variability of the space, coupled with the high temperature and high pressure at high altitude in the flight environment, the pure axial flame data is obviously unable to accurately and comprehensively study the performance of the protective part; the protective part studied in this scheme is not limited to the plate structure, but is considered from more types of structures, coupled with the impact of the flame and the variability of the degree of combustion, the flame will splash to varying degrees at the moment of contact with the protective plate. Therefore, in order to better analyze the installation distance, angle and other more precise position characteristics of the protective part in the later stage, this scheme studies the three-dimensional characteristics of the flame. In this embodiment, the three-dimensional distribution of the temperature characteristics of the flame is studied. In other embodiments, the method of this scheme can also be used to make corresponding modifications to other tooling to obtain the three-dimensional distribution of other data characteristics.

[0041] like Figure 4 As shown in (a), the flame temperature measurement tool 3 includes several sensors 31, and a specially customized high-temperature flow field temperature sensor is selected, which can be used in the engine fuel atomization combustion state of civil aircraft to test the uniformity and stability of the flow field temperature. The entire tool and the sensor housing are made of the same material. The application of the same material can avoid the heat transfer error and poor heat dissipation caused by the thermal properties of different materials. Molybdenum-lanthanum alloy can be selected, which is resistant to high temperature and high Mach. The temperature range is 0 to 2300℃, the Mach number can reach up to 4.5, the accuracy is not less than 3%, and the response time is not less than 30ms, so as to ensure the stability of data acquisition during the test process and reduce the influence of convective heat transfer.

[0042] The flame temperature measuring tool 3 further includes a first mounting plate 32. The shape of the first mounting plate 32 is considered to save materials, facilitate hanging and unified storage. Figure 4 The shape shown in (a) may also be a rectangular plate; a first center hole 33 is opened in the center for the flame to pass through, and the diameter of the first center hole 33 is larger than the maximum diameter of the flame preset in the burn-through test. The diameter difference between the two ensures that all test flames can completely pass through the first center hole 33, avoiding damage to the tooling by the flame, ensuring that the collection process is not disturbed by the collision of the flame, and ensuring the accuracy and integrity of the data.

[0043] Along the outer circumference of the first center hole, a number of sensor mounting holes 34 are radially opened at preset positions. Considering the rationality of data acquisition for different flames and the size of the conventional high-temperature flow field temperature sensor 31, the preset positions are evenly distributed. A sensor mounting hole 34 can be opened every 15-20 degrees in the 360-degree circumference, with a total number of 6-9. Each sensor mounting hole 34 can correspond to the installation of a sensor or not. The actual arrangement is based on the data acquisition requirements. When installing the sensor, a threaded nut 35 can be used for installation. When not installing the sensor, the hole can be blocked with a plug. It is flexible and convenient, and the operation is simple. The relative distance of the sensor deep into the sensor mounting hole 34 is determined by the thread. In this embodiment, the number and position of sensors installed in all sensor mounting holes 34 and the relative distance of the sensors deep into the sensor mounting hole 34 correspond to the circumferential multi-layer and multi-point data acquisition method, which is described in detail in the subsequent method.

[0044] like Figure 5 (a), (b) and (c) show the front view, left view and top view of the flame impact force measurement tool 4, respectively, which also represent the flame heat flux density measurement tool, with only the installed components being different.

[0045] The flame impact force measuring tool 4 and the flame heat flux density measuring tool both include a second mounting plate 42. The shape of the second mounting plate 42 can be a rectangular plate, also taking into account material saving, convenient hanging and unified storage. A second center hole 43 for the flame to pass through is opened in the center of the second mounting plate 42, and the size of the second center hole 43 is consistent with the first center hole 33.

[0046] The second center hole 43 of the flame impact force measurement tool is equipped with at least one pressure sensor 41 and its isolation component to collect the axial distribution of the flame impact force data; in this embodiment, one is installed at the center of the second center hole 43 to obtain the axial distribution of the flame center impact force; in other embodiments, multiple sensors can be set, and the position of the collection end is the same as the circumferential multi-layer multi-point position. The isolation component includes a ceramic sheet 44 and a composite plate 45, which are installed at the front end of the pressure sensor 41 close to the flame to isolate the pressure sensor 41 from the influence of high temperature as much as possible, such as Figure 6 As shown, the composite plate 45 is installed between the ceramic sheet 44 and the surface of the sensor 41.

[0047] The second center hole of the flame heat flux density measuring tool is equipped with at least one heat flux meter to collect the axial distribution of the flame heat flux density data; in this embodiment, one is installed at the center of the second center hole to obtain the axial distribution of the flame center heat flux density; in other embodiments, multiple heat flux meters can be set, and the position of the collection end is the same as the circumferential multi-layer and multi-point position mode, and the position of the heat flux meter corresponds to the position setting of the pressure sensor 41 to establish data connection for subsequent analysis of related data relationships.

[0048] like Figure 7 As shown, the flat test tool 5 includes a third mounting plate 52 for mounting the protective member 51; the shape of the third mounting plate 52 can be a rectangular plate, also in consideration of saving materials, convenient hanging and unified storage; when the protective member 51 is tested, the test protective member 51 is installed in the center of the tooling, and when the protective member 51 is a flat plate, it is fixed around, with 5-7 fixing positions on each side, such as Figure 7 (a) is a front view; when the protective member 51 is a curved plate, it is fixed up and down, and the fixing can be installed by threading, such as Figure 7 (b) As shown in the left view; when conducting an external load test, the influence of the size of the external load 54 on the performance of the protective member 51 is studied. After the test protective member 51 is installed on the upper part of the tooling, different heavy loads are installed on the lower part of the test protective member 51 to conduct corresponding burning tests, such as Figure 7 (c) Front view.

[0049] like Figure 3 (a) and (b) are the front view and right view of the base 2, respectively. The base 2 includes two brackets and at least two parallel crossbars 21 installed between the two brackets. The crossbar 21 has a positioning mark, which can be a positioning mark line or a positioning mark plate for actual positioning to ensure fast and accurate positioning of the tooling; the two crossbars 21 can be pulled out left and right. When performing an external load test, the lower crossbar 21 is pulled out, and only the upper crossbar 21 is retained to avoid the lower crossbar 21 supporting the test guard 51 during the test, which makes it impossible to accurately test the influence of the external load 54, thereby improving the accuracy of the external load 54 test. In this embodiment, the bracket includes a rectangular base plate 22 and a tripod 23 and a vertical rod installed on the base plate 22; the vertical rod and the base plate are movably connected, and the oblique support rod 23 is retractable, thereby adjusting the base angle β to adjust the axial offset angle; the oblique support rod 23 can ensure that the tooling does not tilt backward relative to the base 2; the entire base 2 is heavy, and its own weight can also ensure that the base 2 as a whole does not tilt backward, ensuring that the entire base 2 is stable and does not move during the process of each data acquisition tooling being detachable and switchable and installed on the base 2, thereby maintaining the measurement accuracy of the test data at the same position.

[0050] like Figure 5As shown in (b) and (c) (arrows indicate axial direction), the first mounting plate 32, the second mounting plate 42 and the third mounting plate 52 are all installed with a first special-shaped structure 6 and a second special-shaped structure 7, which are used for detachable installation on the base 2, wherein the first special-shaped structure 6 is adapted to the cross bar 21 to achieve stable suspension and quick removal; the second special-shaped structure 7 can facilitate the placement of tooling with the help of external tools, such as using a clamp to insert the second special-shaped structure 7 to place the tooling. In this embodiment, a first special-shaped structure 6 is installed at the midpoint of the upper part of the first mounting plate 32, the second mounting plate 42 and the third mounting plate 52 facing the flame ejection, and a second special-shaped structure 7 is installed on the side facing the test personnel when in use; the first special-shaped structure 6 and the second special-shaped structure 7 are both opened at the bottom, and the special-shaped structures of the two can be the same. When the cross bar 21 and the external tool clamp are cylindrical, they are divided into an upper arc 61 and a lower arc 62, and the angle of the upper arc 61 is smaller than the angle of the lower arc 62, so that the angle of the lower arc 62 has a large size difference with the cross bar 21 or the external tool clamping part, and the angle of the upper arc 61 has a small size difference with the cross bar 21 or the external tool clamping part, and the shape of the upper and lower parts is different from that of the cross bar 21 and the external tool clamping part. The shape can be designed according to the actual shape of the cross bar 21 or the external tool clamping part to meet the corresponding size difference to achieve the effect of this solution; along the extension direction of the special-shaped shape, the thickness of the two is different, and the thickness of the first special-shaped structure 6 is one-quarter to one-third of the radial length of the mounting plate, ensuring that one first special-shaped structure 6 can also be stably suspended, while avoiding the alignment of the two first special-shaped structures 6 during suspension, thereby increasing the suspension speed; the thickness of the second special-shaped structure 7 is 1-1.2 times the axial thickness of the mounting plate, and the angle difference α between the second special-shaped structure 7 and the mounting plate is 100-120 degrees. The definition of radial and axial here is based on the radial and axial definitions in this system in the installation position state during test use.

[0051] When in use, the external tool can be quickly aligned and passed through the lower arc 62 through the large size difference of the second special-shaped structure 7, and when it moves upward to the upper arc 61 with a small size difference, it can directly face the attached Figure 5 (c) The external tool is rotated counterclockwise at a certain angle to compensate for the angle difference, so that the external tool can stably clamp the tool under the reaction force of its own weight and the angle difference, and can be stably and smoothly moved to the base 2; then, the tool is quickly hung on the crossbar 21 through the large size difference lower arc 62 of the first special-shaped structure 6, and the tool can be stably suspended on the crossbar 21 without displacement through the small size difference upper arc 61 and the self-weight of the tool. After the test is completed, the tool is removed from the base 2 and another tool is switched to be hung. The whole process can also ensure stability and smoothness under the special design of the special-shaped structure, without affecting the positioning of the base 2, and is fast and labor-saving.

[0052] When conducting tests in the prior art, there is always a switching waiting time, such as system adjustment and tooling cooling. Compared with the prior art, this solution is based on the optimization of the method process and improves the tooling structure. The size of the base 2 and the tooling and the hanging position are set according to the center point of the flame ejection. With the special design of the positioning mark and the arc structure, quick disassembly and assembly, accurate positioning, and position consistency can be achieved, ensuring that under the same test conditions, the switching time is reduced as much as possible, and operational and accurate data collection is achieved under a safe and as consistent a thermal environment as possible, ensuring the accuracy of the test results.

[0053] like Figure 2 As shown, the present embodiment also provides an aircraft engine combustion chamber casing burn-through test system, including an aircraft engine combustion chamber casing burn-through test device, a gas generator 1, and a nitrogen supply system, an air supply system, an oxygen system, a kerosene supply system, a water supply system, an automatic gas distribution system, and a measurement and control system connected to the gas generator 1.

[0054] The nitrogen supply system, air supply system, oxygen system, kerosene supply system, water supply system, automatic gas distribution system and measurement and control system mainly supply the fuel, oxidant, cooling water and auxiliary gas required for the operation of the gas generator 1. The preset air and aviation kerosene are transported to the gas generator 1 to organize combustion and form high-temperature combustion gas. The high-temperature combustion gas passes through the throat of the gas generator 1 and sprays flames from its nozzle. The flame is used to perform ablation test of the protective part 51 on the test piece. During the ablation test, the burn-through test device of the combustion chamber casing of the aircraft engine is placed at the front end of the nozzle of the gas generator 1 and used according to the burn-through test method of the combustion chamber casing of the aircraft engine.

[0055] Specifically, the gas generator 1 mainly includes four parts: the head, the body, the throat and the torch igniter 11. The combustion temperature is relatively high. The combustion is organized by the head air + fuel spray atomization method, the torch igniter 11 is radially ignited, and the body and throat interlayer water cooling technical solution. When the gas generator 1 is working, the air and aviation kerosene that meet the flow and pressure requirements are respectively supplied to the corresponding medium inlet of the gas generator 1 by the air supply system and the kerosene supply system. The air enters the air nozzle through the air collecting cavity and enters the combustion chamber. The kerosene enters the kerosene nozzle through the liquid collecting cavity and is sprayed into the combustion chamber after atomization. The air and kerosene are radially ignited and burned in the combustion chamber by the torch igniter 11 to form high-temperature fuel gas that meets the requirements. Cooling water is supplied to the interlayer of the body and throat by the water supply system for cooling to ensure that the gas generator 1 can work reliably for a long time. The main functions of the oxygen supply system are to provide oxygen with set pressure and flow to the torch igniter 11. The main function of the nitrogen supply system is to purge the system before and after ignition. It is divided into igniter oxygen kerosene purge and gas generator 1 kerosene purge, providing nitrogen for kerosene container pressurization and system drive nitrogen. The main function of the air compressor and post-processing is to compress the air in the atmospheric environment to the required pressure, and fill it into the high-pressure air storage tank for storage, providing high-pressure air for the air supply system. The automatic gas distribution supply system realizes full process automation control and is the driving command mechanism for all pneumatic pressure reducing valves to achieve pressure regulation. The test personnel set the required parameters on the control system operation interface according to the test parameter requirements, and adjust the pressure to meet the requirements until the test is completed, which can be operated remotely. The measurement and control system is used to measure various test parameters such as pressure, temperature, flow, etc. of the test equipment, and at the same time complete the automatic control of the working process of the test system, and monitor the working status of the test equipment. It has automatic timing and emergency shutdown functions.

[0056] The use of the system of this solution has many adjustable parameters and can simulate various engine combustion conditions, which is conducive to establishing the relationship between the ablation condition of the protective component 51 and the engine combustion condition, and can reversely judge the cause of the engine combustion from the ablation degree of the protective component 51. In addition, the use of the system makes the test repeatable, which can ensure the consistency of the test conditions and effectively support the subsequent test data relationship analysis.

[0057] According to the above-mentioned aircraft engine combustion chamber casing burn-through test system and device, an aircraft engine combustion chamber casing burn-through test method is provided, such as Figure 1 As shown, the method includes:

[0058] Step S100 is to simulate the combustion environment in which flames from the combustion chamber of an aircraft engine are ejected and the protective element 51 is burned in a simulated flight environment.

[0059] Specifically, it is realized through the burn-through test system of the combustion chamber of an aircraft engine, and the burn-through flame of the casing can be formed by setting the configuration parameters of the system. According to relevant experimental research and FAA regulations, when the combustion chamber casing burns through, if the burn-through aperture is 1 inch (1 inch = 25.4mm), the jet fire flame temperature formed can reach 1648 ℃. This system can be built for reference, which can not only meet the requirements of the existing conventional burn-through test regulations, but also build more combustion environments on this basis. The configuration parameters include nozzle aperture, oil-gas mass ratio, outer air flow rate and combustion chamber internal pressure. Based on the standard test requirements, the nozzle aperture can be set to 12.7mm, 25.4mm and 50.8mm; the oil-gas mass ratio is 0.035-0.050, with a value of 0.005 every time; the outer air flow rate is 0m / s, 20m / s, 50m / s and 70m / s; the combustion chamber internal pressure is 3.5mpa, 4mpa and 4.5mpa.

[0060] Step S200, flame characteristic test:

[0061] S201, within a preset time after the flame is ejected with the determined configuration parameters, each flame data acquisition tool starts to move away from the flame ejection starting point in sequence according to the axial position change speed, and collects corresponding flame data during the axial position change process; wherein each flame data acquisition tool is switched in a detachable installation manner, and at least one data acquisition tool obtains the corresponding flame characteristic data at each axial position in a circumferential multi-layer and multi-point data acquisition manner, and aggregates to obtain a three-dimensional distribution of the corresponding flame characteristic data.

[0062] Specifically, the preset time after the flame is ejected may be 3 minutes, and the ejection mode may be continuous to simulate the flame conditions and combustion conditions under different flight environments.

[0063] like Figure 8 As shown, the direction indicated by the arrow is the axial direction, and the axial position change speed is the speed at which the flame data acquisition tool moves axially away from the starting point of the flame ejection. It can be set to move at a speed of 16-20mm / s. The maximum axial position of the conventional flame data acquisition tool and the starting point of the flame ejection is 800mm. This speed setting can ensure that comprehensive data is collected when multiple data acquisition tools are switched within the preset time.

[0064] The flame characteristic test is carried out using the flame temperature measurement tool 3, the flame heat flux density measurement tool 4 and the flame impact force measurement tool. Any of the three toolings is hung on the base 2, and the flame is ejected through the system according to the configuration parameters. The base 2 starts to move away from the flame ejection starting point according to the axial position change speed with the help of an external control device. During the axial position change process, the corresponding flame characteristic data (flame temperature, flame impact force or flame heat flux density) is obtained through the components on the tooling (temperature sensor, pressure sensor or heat flux meter); the base 2 stops when it moves to the maximum axial position, the tooling completes data acquisition, switches to other flame data acquisition tools, and the test is carried out again according to the above process.

[0065] In this embodiment, the tooling for realizing the circumferential multi-layer multi-point digital sampling method is the flame temperature measurement tooling 3. The circumferential multi-layer multi-point digital sampling method is to determine the circumference with the axial direction determined by the axial distance, and determine at least two circumferential layers with a radius difference with the center point of the flame as the center of the circle. The radius difference and the number of points on each circumferential layer, the point position and the point density are arranged in a way that can reflect the three-dimensional shape of the flame in the flight environment. That is, the number and position of the sensors installed on the first mounting plate 32 of the flame temperature measurement tooling 3 correspond to the number of points, the point position and the point density on the circumferential layer, and the distance between the acquisition end of the sensor and the center hole and the radius difference of the circumferential layer correspond.

[0066] like Figure 4As shown in (b) (simplified diagram, the dotted arrow indicates radial direction, and the curved arrow indicates circumferential direction), in order to better describe the arrangement, the sensor mounting holes 34 are numbered, for example, 9 sensor mounting holes 34 are evenly distributed (9 small hollow circles are shown in the figure), the bottom of the front view is numbered 1, and the numbers are 1-9 clockwise. One flame center point and two circumferential layers are selected, 4 points are selected in the outer circumferential layer, and 4 points are selected in the inner circumferential layer. Therefore, a sensor is installed at position number 4, and the acquisition end (solid circle in the figure) is aligned with the first center hole 33. Hole center, collect the axial data distribution of the flame center temperature; install a sensor at position numbers 1, 2, 6, and 9 respectively, and the distance between each sensor collection end and the hole center of the first center hole 33 is the inner circumferential layer radius; install a sensor at position numbers 3, 5, 7, and 8 respectively, and the distance between each sensor collection end and the hole center of the first center hole 33 is the outer circumferential layer radius; the sensor is installed through a threaded nut assembly, and the distance between the sensor collection end and the hole center of the first center hole 33 is determined by the number of threads or position or mark. In this embodiment, the outer circumferential layer diameter can be 15.7mm, and the inner circumferential layer diameter d2 can be 12.7mm. This number and position arrangement can well collect data and obtain flame three-dimensional data, ensuring the rationality of the amount of collected data and the accuracy of being able to characterize the three-dimensional flame characteristics under the flight environment. Of course, the specific position number and radius can be arranged according to actual needs. In this embodiment, when the test is carried out at different axial distances, the point positions on different circumferential layers are static and remain motionless.

[0067] S202, aggregating the data acquired by all flame data acquisition tools to obtain the flame characteristic data under the configuration parameters and axial position change speed; specifically, aggregating to obtain the corresponding flame characteristic data distribution collected by the tool under the flame characteristic test condition of the configuration parameters, that is, the flame temperature distribution, the flame impact force distribution and the flame heat flux density distribution.

[0068] S203, adjusting the configuration parameters and the axial position change speed, and repeating the above steps to obtain corresponding flame characteristic data.

[0069] Step S300, a burn-through test of the protective component is performed. After the three fixtures complete the flame characteristic test, the flat test fixture 5 with the protective component 51 installed is switched to perform an ablation test and an external load test of the protective component 51 under the flame characteristic test condition with the configuration parameters.

[0070] S301, based on the flame characteristic data obtained from the configuration parameters in S200 and the axial position change speed, determine the axial distance between the flat test fixture and the flame ejection starting point, and obtain the historical experience axial distance to obtain an axial distance set.

[0071] Specifically, the axial distance between the flat test fixture and the starting point of flame ejection is determined by using the flame characteristic data, and the axial distance is determined by the axial position corresponding to the first three data that clearly characterize the corresponding flame characteristics in the corresponding flame characteristic data obtained by each flame data acquisition fixture. That is, the temperature, impact force, and heat flux data are sorted according to the flame temperature distribution, flame impact force distribution, and flame heat flux density distribution, and the axial positions corresponding to the three data are taken from high to low. For example, the axial position corresponding to the highest heat flux density is 300mm away from the starting point of flame ejection of the fixture, and 300mm is taken as an axial distance. By analogy, an axial distance set is obtained.

[0072] For the axial distance, an axial distance spacing threshold is also set, such as 50-80 mm. When the spacing between two adjacent axial distances in the axial distance set is smaller than the threshold, an axial distance is determined by the average value of the two axial distances.

[0073] Conventional historical experience axial distances can be selected from at least 5, 127mm, 254mm, 381mm, 508mm and 762mm. However, there are limitations to using only the historical experience axial distance for burn-through tests, which is divorced from the current flame environment. Therefore, in order to improve the accuracy of the burn-through test of protective parts and make a more comprehensive evaluation of protective parts, this scheme determines the axial distance based on the flame characteristic data obtained, and then selects the axial distance based on the threshold value in combination with the historical experience axial distance. This not only conforms to the authenticity of the current test, but also combines historical experience. At the same time, it can also associate the axial distance of the burn-through test of protective parts with the current flame characteristic data, which is conducive to relationship analysis.

[0074] S302, within a preset time after the flame is sprayed with the configuration parameters in S200, the non-curved protective parts are installed in sequence using the flat test fixture, and the burn-through test is performed at all axial distances in the axial distance set to obtain the protective part ablation data corresponding to the configuration parameters and the axial distance set.

[0075] Specifically, the ablation data of the protective element 51 is obtained by measuring the ablated protective element 51 and recording the ablation time to obtain the ablation data of the protective element 51, including the deformation during the fire exposure time and the effective protection time, and also the deformation pressure.

[0076] S303, based on the protective part ablation data obtained from the protective part without curvature in S302, determine the axial offset angle between the flat test fixture and the flame ejection starting point and the basic parameter curvature radius of the protective part, adjust the axial offset angle and the curvature radius of the protective part, and repeat S302 to obtain the corresponding protective part ablation data.

[0077] Specifically, a 3D scanner is used to scan the deformed area of ​​the non-curvature protective part after the test to obtain a point cloud data set, and the data is imported into a 3D surface morphology analysis software such as MountainsMap / Mountainslab to obtain a curvature radius data set and an inclination angle data set of the deformed area. The maximum, minimum, and average values ​​in the curvature radius data set are selected to form a test curvature radius set, and the maximum, minimum, and average values ​​in the inclination angle data of the lower half of the deformation area are selected to form an axial offset angle set for the test.

[0078] Based on the experimental curvature radius set and experimental axial offset angle set obtained based on the current flame characteristic data, the curvature radius and axial offset angle determined by conventional experience can also be added to increase the test and improve the support for relationship analysis. There are 4 optional curvature radii, 5in, 10in, 15in and 20in; there are at least 4 optional axial offset angles, 30°, 45°, 60° and 90°. At the same time, it can also be filtered by threshold in the same way as the axial distance filtering method.

[0079] At each axial distance, the axial offset angle of the protective member 51 is changed through the base 2 to obtain the corresponding ablation data of the protective member 51 at the axial distance and the axial offset angle; the flat plate test fixture 5 replaces and installs the protective members 51 with different curvature radii to perform the test according to the above process to obtain the corresponding ablation data of the protective members 51 with different curvature radii.

[0080] S304, adjusting other basic parameters of the protective element, such as material thickness and material type, and repeating the above steps to obtain corresponding protective element ablation data.

[0081] Step S400, based on the flame characteristic data and the shield ablation data, the relationship between the flame configuration parameters, the shield basic parameters, the axial distance, and the axial offset angle is analyzed. After a large amount of test data is obtained, comparative analysis can be performed through experience, or by building a model algorithm, where the model algorithm can be implemented according to the prior art.

[0082] The burn-through test method, system and device of the combustion chamber casing of an aircraft engine provided in this embodiment can reduce environmental differences and ensure environmental consistency in a continuous thermal space that can reflect the combustion environment of an aircraft engine through the ingenious setting of tooling structures and data collection points. The test time can be shortened by fast switching, and three-dimensional flame characteristic data and protective component ablation data can be obtained more quickly, thereby reducing the error between test data and authenticity, and obtaining more accurate test results based on the limited amount of data obtained. At the same time, the axial offset angle and curvature radius change of the protective component are determined by the actual test data, which is conducive to studying the correlation between the data, so as to study the relationship between the burn-through degree of the protective component and the combustion condition of the engine, and to obtain a more complete judgment of the performance of the protective component, which promotes the precise installation position of the more three-dimensional new protective component.

[0083] Embodiment 2

[0084] What is different from the first embodiment is that, along the circumference of the first center hole, the first mounting plate 32 belonging to the flame temperature measuring tool 3 is provided with a plurality of first corresponding hole positions, and holes corresponding to the first corresponding hole positions are opened on the first special-shaped structure 6 belonging to the first mounting plate. By aligning the holes with different first corresponding hole positions respectively, the first special-shaped structure is installed on the first mounting plate by rotating the first angle, so that different angles of the first mounting plate 32 in the circumferential direction can be achieved to form a plurality of rotation points.

[0085] The specific use process is that in S40, when the test is carried out at different axial distances, the position of the points on different circumferential layers is dynamically rotated and changed. That is, the first mounting plate 32 collects data under the same test condition at the initial point and several rotation points respectively; the rotation test is that the first mounting plate 32 is the initial point, and the initial point test is first carried out under the test condition; then the test under the test condition is carried out in turn at different rotation points, and the rotation point can be adjusted clockwise along the circumference. Each rotation point can be spaced 20-40 degrees apart, and the number of rotations is 3-5 times. By setting a reasonable rotation angle and number of times, it is possible to obtain enough data on the basis of ensuring rapid testing, and aggregate to obtain a more complete three-dimensional distribution of the flame.

[0086] Embodiment 3

[0087] Different from the first and second embodiments, the base 2 also includes a guide rail, which is connected to a drive device and a control system, and the bracket of the base 2 is installed on the guide rail. The control system sends a signal to the drive device to realize the axial movement and angle change of the guide rail, so as to achieve the accuracy of the axial distance and axial offset angle change of the entire base 2. Compared with the telescopic adjustment of the oblique support rod 23, this solution increases the speed of position data change during the entire test process through the automatic adjustment switching mode of the base 2, further reduces the test time, and improves the accuracy of the results.

[0088] The above is only an embodiment of the present invention. The common sense such as the known specific structure and characteristics in the scheme is not described in detail here. The ordinary technicians in the relevant field are aware of all the common technical knowledge in the technical field of the invention before the application date or priority date, can obtain all the existing technologies in the field, and have the ability to apply the conventional experimental means before that date. The ordinary technicians in the relevant field can improve and implement the scheme in combination with their own abilities under the enlightenment given by this application. Some typical known structures or known methods should not become obstacles for the ordinary technicians in the relevant field to implement this application. It should be pointed out that for the technicians in this field, without departing from the structure of the present invention, several deformations and improvements can be made, which should also be regarded as the protection scope of the present invention, and these will not affect the effect of the implementation of the present invention and the practicality of the patent.

Claims

1. A burn-through test method for a combustion chamber casing of an aircraft engine, characterized in that: The method comprises: Step S100, simulating a combustion environment in which flames are ejected from a combustion chamber casing of an aircraft engine and ablates protective components in a flight environment; Step S200, flame characteristic test: S201, within a preset time after the flame is ejected with the determined configuration parameters, a plurality of flame data acquisition tools are used to sequentially move away from the flame ejection starting point according to the axial position change speed, and corresponding flame data are collected during the axial position change process; wherein each flame data acquisition tool is switched in a detachable installation manner, and at least one data acquisition tool obtains corresponding flame characteristic data at each axial position in a circumferential multi-layer and multi-point data acquisition manner, and aggregates to obtain a three-dimensional distribution of the corresponding flame characteristic data; S202, aggregating data acquired by all flame data acquisition tools to obtain flame characteristic data under the configuration parameters and axial position change speed; S203, adjusting the configuration parameters and the axial position change speed, and repeating the above steps to obtain corresponding flame characteristic data; Step S300, performing a burn-through test of the protective component: S301, determining the axial distance between the flat test fixture and the flame ejection starting point based on the flame characteristic data obtained from the configuration parameters and the axial position change speed in S200, and obtaining the historical experience axial distance to obtain an axial distance set; S302, within a preset time after the flame is sprayed with the configuration parameters in S200, the non-curved protective parts are installed in sequence using the flat test fixture, and a burn-through test is performed at all axial distances in the axial distance set to obtain the protective part ablation data corresponding to the configuration parameters and the axial distance set; S303, based on the shield ablation data obtained from the shield without curvature in S302, determine the axial offset angle between the flat test fixture and the flame ejection starting point and the basic parameter curvature radius of the shield, adjust the axial offset angle and the curvature radius of the shield, and repeat S302 to obtain the corresponding shield ablation data; S304, adjusting other basic parameters of the protective element, such as material thickness and material type, and repeating the above steps to obtain corresponding protective element ablation data; Step S400, based on the flame characteristic data and the shield ablation data, analyzing and obtaining the relationship between the flame configuration parameters, the shield basic parameters, the axial distance, and the axial offset angle; The flame data acquisition tools include a flame temperature measurement tool, a flame heat flux density measurement tool and a flame impact force measurement tool; The flame data includes flame temperature, flame heat flux density and flame impact force.

2. The burn-through test method of the combustion chamber casing of an aircraft engine according to claim 1, characterized in that: In S201, the circumferential multi-layer and multi-point number sampling method is to determine the circumference with the axial direction determined by the axial position, and determine at least two circumferential layers with a radius difference with the flame center point as the center of the circle. The radius difference and the number of points on each circumferential layer, the point position and the point density are arranged in a way that can reflect the three-dimensional shape of the flame in the flight environment.

3. The burn-through test method of the combustion chamber casing of an aircraft engine according to claim 2, characterized in that: When testing at different axial positions, the positions of points on different circumferential layers are either kept statically or changed dynamically.

4. The burn-through test method of the combustion chamber casing of an aircraft engine according to claim 1, characterized in that: In S301, the axial distance between the flat test fixture and the flame ejection starting point is determined based on the flame characteristic data obtained based on the configuration parameters in S200 and the axial position change rate, and the axial distance is determined based on the axial position corresponding to the first three data that clearly characterize the corresponding flame characteristics in the corresponding flame characteristic data obtained by each flame data acquisition fixture.

5. The burn-through test method of the combustion chamber casing of an aircraft engine according to claim 4, characterized in that: An axial distance spacing threshold is set. When the spacing between two adjacent axial distances in the axial distance set is smaller than the threshold, an axial distance is determined by the average value of the two axial distances.

6. The burn-through test method of the combustion chamber casing of an aircraft engine according to claim 1, characterized in that: The flame characteristic data include flame temperature distribution, flame heat flux density distribution and flame impact force distribution; the protection component ablation data include deformation during the fire exposure time and effective protection time.

7. The burn-through test method of the combustion chamber casing of an aircraft engine according to claim 6, characterized in that: In S303, the axial offset angle between the flat test fixture and the flame ejection starting point and the basic parameter curvature radius of the protective part are determined based on the protective part ablation data obtained from the non-curvature protective part in S302. A three-dimensional scanner is used to scan the deformation area of ​​the non-curvature protective part after the test to obtain a point cloud data set, which is imported into the three-dimensional surface morphology analysis software to obtain the curvature radius data set and the inclination angle data set of the deformation area. The maximum, minimum and average values ​​in the curvature radius data set are selected to form a curvature radius set for the test. The maximum, minimum and average values ​​in the inclination angle data of the lower half of the deformation area are selected to form an axial offset angle set for the test.

8. An aircraft engine combustion chamber casing burn-through test device, which implements the method for using the data acquisition tool in the aircraft engine combustion chamber casing burn-through test method as described in any one of claims 1 to 7, characterized in that: It includes a base and a flame temperature measuring tool, a flame heat flux density measuring tool, a flame impact force measuring tool and a flat plate test tool that can be detachably mounted on the base; The flame temperature measuring tool comprises a first mounting plate and a plurality of sensors; the first mounting plate is provided with a first central hole for the flame to pass through; a plurality of sensor mounting holes are radially provided at preset positions along the outer circumference of the central hole, wherein the number and positions of the sensors installed in all the sensor mounting holes and the relative distances of the sensors into the sensor mounting holes correspond to the circumferential multi-layer and multi-point digital sampling method; The flame heat flux density measurement tool and the flame impact force measurement tool both include a second mounting plate; the second mounting plate is provided with a second center hole for the flame to pass through; at least one heat flux meter is installed in the second center hole of the flame heat flux density measurement tool; at least one pressure sensor and its isolation component are installed in the second center hole of the flame impact force measurement tool; the flat plate test tool includes a third mounting plate for installing a protective part; wherein the first mounting plate, the second mounting plate and the third mounting plate are all provided with a first special-shaped structure and a second special-shaped structure for detachable installation.

9. The aircraft engine combustion chamber casing burn-through test device according to claim 8, characterized in that: Along the circumference of the first center hole, a plurality of first corresponding hole positions are opened on the first mounting plate to which the flame temperature measuring tool belongs, and holes corresponding to the first corresponding hole positions are opened on the first special-shaped structure to which the first mounting plate belongs. The holes are aligned with different first corresponding hole positions respectively, and the first special-shaped structure is rotated and installed on the first mounting plate.

10. Aircraft engine combustion chamber casing burn-through test system, characterized in that: It includes a gas generator and a nitrogen supply system, an air supply system, an oxygen system, a kerosene supply system, a water supply system, an automatic gas distribution system and a measurement and control system connected to the gas generator, which are used to simulate the combustion environment of the aircraft engine combustion chamber casing flame ejection and ablation of protective parts in a flight environment, forming different casing burn-through flames and their corresponding configuration parameters; it also includes an aircraft engine combustion chamber casing burn-through test device as described in any one of claims 8-9, which is placed at the front end of the gas generator flame ejection and is used to collect flame characteristic data and protective part ablation data.

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