Thermal shock test device, method and equipment for turbine blade under near-service condition

Abstract

The invention relates to the technical field of aero-engines, and discloses a thermal shock test device, method and equipment for turbine blades under near-service conditions, and the device comprises a fuel oil pressurization module which provides fuel oil for a combustion module; the combustion module is used for combusting air and fuel oil to generate high-temperature fuel gas meeting preset conditions; the temperature acquisition module is used for acquiring the real-time temperature of the high-temperature fuel gas; and the controller is used for acquiring the real-time temperature and controlling the amount of the fuel oil entering the combustion module according to the real-time temperature, so that the real-time temperature changes according to a preset temperature rise curve. The combustion module is designed to combust aviation kerosene, it is ensured that the loaded condition of the tested turbine blade in the thermal shock test process is basically kept consistent with the loaded condition in the engine in the service state, the fuel pressurization module is controlled by the controller to adjust the amount of fuel entering the combustion module, and the accuracy of the test temperature is ensured; the environment of the thermal shock test is closer to engineering practice, and the service life of the turbine blade and the coating at high temperature is accurately evaluated.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engines, and particularly to a thermal shock test device, method and equipment for turbine blades under near-service conditions. Background Art

[0002] With the development of aero-engines towards high thrust-to-weight ratios, the gas temperature in front of the turbine of the engine has been continuously increasing, and the working environment of the turbine blades has become increasingly harsh. The impact of high-temperature gas will have a certain impact on the strength and life of the turbine blades. Therefore, mastering the thermal cycle life of turbine blades under high-temperature gas impact is of great significance for evaluating the safe life of turbine blades during service. At present, the thermal shock test is mainly used to study the thermal cycle resistance of materials. However, the existing thermal shock test methods are more targeted at material specimens or test bars. On the one hand, the influence of the blade structure is not considered, and at the same time, the actual temperature load of the blade under the engine service environment is rarely combined, resulting in a certain gap between the test results and the actual service environment, which has a certain impact on evaluating the thermal shock resistance of the blades.

[0003] Due to the long-term impact of high-temperature gas on the working blades of gas turbines, common failure forms include thermal fatigue damage, surface coating peeling, high-temperature oxidation, etc. The thermal shock test is mainly used to simulate the load characteristics of materials or parts under thermal cycling and gas impact, and to evaluate the thermal fatigue resistance and high-temperature oxidation ability of materials or parts. For the existing thermal shock test devices or methods, the high-temperature environment is generally provided by thermocouple heating, without fully considering the gas environment characteristics after the combustion of aviation kerosene inside the aero-engine, and the test objects are mostly material-level specimens, and there are certain differences between the structural forms of the test pieces and the real blades; in addition, some thermal shock test devices or methods only set the load spectrum according to the highest and lowest temperatures of the hot-end components of the engine, with only two temperature points alternating. However, in actual operation, the engine generally has multiple states such as ground idle, air idle, maximum takeoff, maximum continuous, and cruise. The holding time of each different state and the corresponding temperature characteristics are different. Only considering the lowest and highest temperatures does not conform to the actual engineering situation, resulting in a large deviation in the evaluation of the thermal shock resistance of turbine blades or coatings under the service environment. Summary of the Invention

[0004] In view of this, the present invention provides a thermal shock test device, method and equipment for turbine blades under near-service conditions to solve the problem of large deviation in the evaluation of the thermal shock resistance of turbine blades or coatings under the service environment due to the thermal shock test conditions not conforming to the actual engineering situation.

[0005] In a first aspect, the present invention provides a thermal shock test device for turbine blades under near-service conditions, and the device includes: a fuel boosting module, a controller, a combustion module, and a temperature acquisition module, wherein,

[0006] A fuel boosting module, whose outlet end is connected to the fuel inlet end of the combustion module, is used to supply fuel to the combustion module;

[0007] A combustion module, whose air inlet end admits air, uses air and fuel for combustion to generate high-temperature gas meeting preset conditions, and the high-temperature gas acts on the test turbine blade for a thermal shock test;

[0008] A temperature acquisition module, which is used to acquire the real-time temperature of the high-temperature gas at the outlet end of the combustion module;

[0009] A controller is used to obtain the real-time temperature of the high-temperature gas and control the fuel boosting module according to the real-time temperature to change the fuel quantity entering the combustion module, so that the real-time temperature changes according to a preset heating curve.

[0010] The thermal shock test device for turbine blades under near-service conditions provided by the present invention designs the combustion module to burn aviation kerosene, ensuring that the loading conditions of the test turbine blade during the thermal shock test are basically the same as those in the engine during service. The real-time temperature reaching the test turbine blade is obtained by using the temperature acquisition module, and the fuel boosting module is adjusted according to the real-time temperature to ensure the accuracy of the test temperature, so as to make the environment of the thermal shock test closer to the engineering actual situation and accurately evaluate the service life of the turbine blade and coating at high temperatures.

[0011] In an optional embodiment, the fuel boosting module includes: a fuel tank, a first main fuel valve, and a first sub-fuel valve, and the combustion module includes: a primary combustion chamber. Among them,

[0012] The fuel tank, whose outlet end is respectively connected to the first end of the first main fuel valve and the first end of the first sub-fuel valve, is used to supply fuel to the combustion module;

[0013] The first main fuel valve, whose second end is connected to the fuel inlet end of the primary combustion chamber, is used to change the opening degree according to the first instruction of the controller and adjust the fuel quantity entering the primary combustion chamber. The first instruction is the opening degree change amount of the first main fuel valve obtained by the controller through data processing based on the real-time temperature and the preset heating curve;

[0014] The first sub-fuel valve is connected in parallel with the first main fuel valve and is used to assist the first main fuel valve in controlling the fuel quantity entering the primary combustion chamber;

[0015] The primary combustion chamber, whose air inlet end admits air, uses air and fuel for combustion to generate first high-temperature gas meeting the first preset conditions, and the first high-temperature gas acts on the test turbine blade for a thermal shock test.

[0016] In an alternative embodiment, the fuel boosting module further includes: a second main fuel valve and a second auxiliary fuel valve. The combustion module further includes a secondary combustion chamber, wherein,

[0017] The second main fuel valve, its first end is connected to the outlet end of the fuel tank, and its second end is connected to the fuel inlet end of the secondary combustion chamber, and is used to change the opening degree according to the second instruction of the controller to adjust the fuel quantity entering the secondary combustion chamber. The second instruction is the opening degree change amount of the second main fuel valve obtained by the controller through data processing according to the real-time temperature and the preset temperature rising curve;

[0018] The second auxiliary fuel valve, which is connected in parallel with the second main fuel valve, and is used to assist the second main fuel valve to control the fuel quantity entering the secondary combustion chamber;

[0019] The secondary combustion chamber, its air inlet end is connected to the outlet end of the primary combustion chamber and the gas pipeline, and air and the first high-temperature gas are introduced. The air and fuel are used for combustion to heat the first high-temperature gas, and the second high-temperature gas that meets the second preset condition is generated. The second high-temperature gas acts on the test turbine blade for thermal shock test.

[0020] In an alternative embodiment, the temperature acquisition module includes: a plurality of thermocouples. Each thermocouple includes a plurality of temperature measurement points. The plurality of temperature measurement points of the plurality of thermocouples are arranged in a concentric circle distribution manner on the gas outlet cross-section of the combustion module.

[0021] The thermal shock test device for turbine blades under near-service conditions provided by the present invention ensures the full combustion of aviation kerosene and reasonably regulates the temperature uniformity at the outlet of the secondary combustion chamber through single-stage combustion chamber separate debugging and two-stage combustion chamber joint debugging. By collecting the real-time temperatures at different positions at the outlet of the secondary combustion chamber and feeding them back to the controller for joint debugging, the accuracy of temperature detection is improved. The two main fuel valves are mainly controlled automatically, and at the same time, auxiliary fuel valves are added to regulate the temperatures of the two-stage combustion chambers. When the automatic control system is difficult to achieve the entire range of the set temperature curve, the auxiliary fuel valves are manually adjusted to increase the gas pressure of the two-stage combustion chambers, thereby raising the temperature upper limit of the combustion chamber and the test article and expanding the allowable temperature range of the thermal shock test.

[0022] In a second aspect, a thermal shock test method for turbine blades under near-service conditions provided by the present invention. The thermal shock test method is applied to the thermal shock test device in any one of the first aspect. The method includes:

[0023] Obtain the real-time temperature of the test turbine blade, and compare the real-time temperature with the set temperature corresponding in the preset temperature rising curve;

[0024] Adjust the fuel quantity entering the combustion module according to the comparison result, change the real-time temperature of the tested turbine blade, so that the real-time temperature meets the preset conditions, and the temperature change curve of the tested turbine blade coincides with the preset heating curve.

[0025] The thermal shock test method for turbine blades under near-service conditions provided by the present invention ensures that the loading conditions of the tested turbine blades during the thermal shock test are basically the same as those inside the engine during the service state by designing a two-stage combustion chamber to burn aviation kerosene, and uses the single-stage adjustment and two-stage joint adjustment of the two-stage combustion chamber to ensure the accuracy of the test temperature, so as to make the environment of the thermal shock test closer to the actual engineering situation and accurately evaluate the service life of turbine blades and coatings at high temperatures.

[0026] In an optional implementation manner, obtaining the real-time temperature of the tested turbine blade and comparing the real-time temperature with the set temperature corresponding in the preset heating curve includes:

[0027] Obtain the real-time temperature of multiple temperature measurement points on the cross-section of the gas outlet of the combustion module;

[0028] Calculate the real-time temperature of the high-temperature gas as the real-time temperature of the tested turbine blade according to the real-time temperature of multiple temperature measurement points;

[0029] Analyze the temperature field of the turbine blades in multiple service states of the engine to obtain the preset heating curve of the turbine blades;

[0030] Compare the real-time temperature of the tested turbine blade with the set temperature corresponding in the preset heating curve at the same sampling moment to obtain the comparison result.

[0031] In an optional implementation manner, adjusting the fuel quantity entering the combustion module according to the comparison result includes:

[0032] When the real-time temperature is less than the set temperature corresponding in the preset heating curve and the deviation between the real-time temperature and the set temperature corresponding in the preset heating curve is greater than the upper limit of the preset deviation range, increase the fuel quantity entering the combustion module;

[0033] When the real-time temperature is greater than the set temperature corresponding in the preset heating curve and the deviation between the real-time temperature and the set temperature corresponding in the preset heating curve is greater than the upper limit of the preset deviation range, reduce the fuel quantity entering the combustion module.

[0034] In an optional implementation manner, increasing the fuel quantity entering the combustion module includes:

[0035] Increase the opening degree of the main fuel valve in the fuel boosting module, and the fuel boosting module is used to provide fuel for the combustion module;

[0036] When the opening degree of the main fuel valve is at its maximum, if the real-time temperature is lower than the set temperature corresponding to the preset temperature rise curve, and the deviation between the real-time temperature and the set temperature corresponding to the preset temperature rise curve is greater than the upper limit of the preset deviation range, then increase the opening degree of the auxiliary fuel valve in the fuel boosting module until the real-time temperature meets the preset conditions.

[0037] The thermal shock test method for turbine blades under near-service conditions provided by the present invention starts from the actual engineering requirements, aims at the service environment state of turbine working blades, fully combines the actual service loads of the engine, obtains the temperature changes of turbine blades under different service states through finite element analysis, and compiles a test load spectrum equivalent to the actual service environment; designs a two-stage combustion chamber, an intake section, an exhaust section, etc. to simulate the engine structure layout, provides a high-temperature gas environment for the turbine blades to work by burning aviation kerosene, and truly simulates the service environment characteristics of the blades; arranges multiple temperature measurement points evenly in a concentric circle layout, combines the advantages of the automatic control system for temperature debugging, and ensures the accuracy and uniformity of the test temperature through single-stage regulation and two-stage joint regulation of the two-stage combustion chamber. Compared with the existing related technologies of thermal shock tests, the present invention is closer to the actual engineering situation and is of great significance for accurately evaluating the service life of turbine blades and coatings at high temperatures.

[0038] In a third aspect, the present invention provides a computer device, including: a memory and a processor, which are communicatively connected to each other. The memory stores computer instructions, and the processor executes the computer instructions to execute the method according to the first aspect or any corresponding embodiment thereof.

[0039] In a fourth aspect, the present invention provides a computer-readable storage medium, on which computer instructions are stored, and the computer instructions are used to cause a computer to execute the method according to the first aspect or any corresponding embodiment thereof. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] In order to more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required to be used in the description of the specific embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0041] Figure 1 is a schematic structural diagram of a thermal shock test device for turbine blades under near-service conditions according to an embodiment of the present invention;

[0042] Figure 2 is a schematic structural diagram of another thermal shock test device for turbine blades under near-service conditions according to an embodiment of the present invention;

[0043] Figure 3 It is a schematic structural diagram of another thermal shock test device for turbine blades under near-service conditions according to an embodiment of the present invention;

[0044] Figure 4 It is a schematic structural diagram of a temperature acquisition unit in a thermal shock test device for turbine blades under near-service conditions according to an embodiment of the present invention;

[0045] Figure 5 It is a schematic flow diagram of a thermal shock test method for turbine blades under near-service conditions according to an embodiment of the present invention;

[0046] Figure 6 It is a schematic diagram of an engine test run spectrum in a thermal shock test method for turbine blades under near-service conditions according to an embodiment of the present invention;

[0047] Figure 7 It is a schematic diagram of a preset temperature rise curve in a thermal shock test method for turbine blades under near-service conditions according to an embodiment of the present invention;

[0048] Figure 8 It is a schematic hardware structure diagram of a computer device according to an embodiment of the present invention. Specific embodiments

[0049] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts fall within the protection scope of the present invention.

[0050] The embodiments of the present invention provide a thermal shock test method, device, and equipment for turbine blades under near-service conditions. By burning aviation kerosene to provide a high-temperature gas environment for the operation of the turbine blades, and using single-stage regulation and two-stage combined regulation of the two-stage combustion chamber to ensure the accuracy of the test temperature, so as to make the environment of the thermal shock test closer to the actual engineering situation and accurately evaluate the service life of the turbine blades and coatings at high temperatures.

[0051] In this embodiment, a thermal shock test device for turbine blades under near-service conditions is provided. Figure 1 It is a schematic diagram of a thermal shock test device for turbine blades under near-service conditions according to an embodiment of the present invention. As Figure 1 shown, the thermal shock device includes: a fuel boosting module, a controller, a combustion module, and a temperature acquisition module.

[0052] As Figure 1As shown, a fuel boosting module, whose outlet end is connected to the fuel inlet end of the combustion module, is used to supply fuel to the combustion module.

[0053] Specifically, fuel is input at the inlet of the fuel boosting module, and its outlet is connected to the fuel inlet end of the combustion module. The fuel boosting module has a valve, which is used to change the fuel pressure and flow rate and adjust the fuel quantity entering the combustion module. The fuel can be aviation kerosene the same as that in actual engineering, only as an example, but not limited thereto.

[0054] As Figure 1 shown, a combustion module, whose air inlet end admits air, uses the air and fuel for combustion to generate high-temperature gas meeting preset conditions, and the high-temperature gas acts on the tested turbine blade for a thermal shock test.

[0055] Specifically, air enters the combustion module through the air inlet section on the gas pipeline, burns with the fuel entering the combustion module to generate high-temperature gas. The high-temperature gas acts on the tested turbine blade through the gas pipeline, and then is discharged through the outlet section on the gas pipeline.

[0056] The air inlet section (for inputting air) and the exhaust section (for discharging the high-temperature gas after combustion) components are designed with reference to the overall structural layout of the engine, and are used to simulate the gas movement state of the high-temperature gas inside the engine.

[0057] As Figure 1 shown, a temperature acquisition module is used to acquire the real-time temperature of the high-temperature gas at the outlet end of the combustion module.

[0058] Specifically, at the outlet position of the combustion module, a temperature acquisition unit is installed to acquire the real-time temperature of the high-temperature gas generated and output by the combustion module. The real-time temperature of the high-temperature gas is transmitted to the controller as the real-time temperature acting on the tested turbine blade, so as to grasp the temperature change of the tested turbine blade in real time.

[0059] As Figure 1 shown, a controller is used to acquire the real-time temperature of the high-temperature gas, and control the fuel boosting module according to the real-time temperature to change the fuel quantity entering the combustion module, so that the real-time temperature changes according to a preset heating curve.

[0060] Specifically, the real-time temperature of the high-temperature gas is equivalent to the real-time temperature when the high-temperature gas reaches the test turbine blade and acts on it. Therefore, the controller controls the fuel quantity entering the combustion module by obtaining the real-time temperature of the high-temperature gas and according to the temperature relationship corresponding in the preset temperature rise curve, thereby changing the real-time temperature of the high-temperature gas, making the real-time temperature of the high-temperature gas generated after combustion equal to the set temperature corresponding in the preset temperature rise curve, and changing according to the preset temperature rise curve. The thermal shock test device includes multiple parts. This embodiment only involves the fuel boosting module, the controller, the combustion module, and the temperature acquisition module. The remaining parts are set according to actual needs and will not be elaborated here.

[0061] In order to simulate the actual working environment of the turbine blade as much as possible, in this embodiment, the thermal shock test device can design two-stage combustion chambers and select the same aviation kerosene as that in actual service as the carrier for creating the gas environment, so that the high-temperature gas environment is basically the same as the actual working environment of the turbine blade.

[0062] The thermal shock test device for turbine blades under near-service conditions provided in this embodiment burns aviation kerosene through the combustion module to ensure that the loading conditions of the test turbine blade during the thermal shock test are basically the same as those inside the engine during the service state. The real-time temperature reaching the test turbine blade is obtained by using the temperature acquisition module, and the fuel boosting module is adjusted according to the real-time temperature to ensure the accuracy of the test temperature, so as to achieve the effect that the environment of the thermal shock test is closer to the engineering actual situation and accurately evaluate the service life of the turbine blade and the coating at high temperatures.

[0063] In some optional implementation manners, as Figure 2 shown, the fuel boosting module includes: a fuel tank, a first main fuel valve, and a first auxiliary fuel valve, and the combustion module includes: a primary combustion chamber.

[0064] As Figure 2 shown, the fuel tank, the outlet end of which is respectively connected to the first end of the first main fuel valve and the first end of the first auxiliary fuel valve, is used to provide fuel for the combustion module.

[0065] Specifically, the fuel tank is used to store fuel, and its structure is not limited here.

[0066] As Figure 2 shown, the first main fuel valve, the second end of which is connected to the fuel inlet end of the primary combustion chamber, is used to change the opening degree according to the first instruction of the controller and adjust the fuel quantity entering the primary combustion chamber. The first instruction is the opening degree change amount of the first main fuel valve obtained by the controller through data processing based on the real-time temperature and the preset temperature rise curve.

[0067] Specifically, the temperature of the first high-temperature gas generated in the primary combustion chamber is mainly controlled by the first main fuel valve. This valve adopts a negative feedback automatic control mode. After determining the preset temperature rise curve according to the actual situation, when the temperature of the first high-temperature gas is lower than the corresponding temperature in the temperature rise curve, the opening of the first main fuel valve is increased, the fuel quantity entering the primary combustion chamber increases, the combustion temperature in the primary combustion chamber rises, and the temperatures of the first high-temperature gas and the test turbine blade also increase accordingly. When the temperature of the first high-temperature gas is higher than the corresponding temperature in the temperature rise curve, the opening of the first main fuel valve is decreased, the fuel quantity entering the primary combustion chamber decreases, the combustion temperature in the primary combustion chamber drops, and the temperatures of the first high-temperature gas and the test turbine blade also decrease accordingly. Through the negative feedback automatic control of the first main fuel valve, the temperatures of the first high-temperature gas and the test turbine blade change according to the preset temperature rise curve.

[0068] During the whole process, obtaining the real-time temperature of the first high-temperature gas, comparing the real-time temperature with the preset temperature rise curve, and obtaining the opening change amount of the main fuel valve according to the comparison result are all processed and completed by the controller.

[0069] As Figure 2 shown, the first auxiliary fuel valve is connected in parallel with the first main fuel valve and is used to assist the first main fuel valve in controlling the fuel quantity entering the primary combustion chamber.

[0070] Specifically, the first main fuel valve can be a main electro-hydraulic proportional valve, and the first auxiliary fuel valve can be an auxiliary electro-hydraulic proportional valve. The control ends of the first main fuel valve and the first auxiliary fuel valve are respectively connected to the controller, and their valve openings are changed according to the control instructions of the controller. The two fuel valves are connected in parallel to jointly control the fuel quantity entering the primary combustion chamber from the fuel tank. Different fuel quantities in the primary combustion chamber result in different temperatures of the high-temperature gas generated by combustion. The more the fuel quantity, the higher the temperature of the high-temperature gas.

[0071] The first auxiliary fuel valve is connected in parallel with the first main fuel valve and controls two fuel oil circuits in the primary combustion chamber respectively. The automatic control system of the first main fuel valve may be difficult to achieve the entire range of the preset temperature curve. For example, when the opening of the first main fuel valve is at its maximum, the temperature of the test turbine blade is still difficult to reach the corresponding temperature in the preset temperature curve. At this time, the opening of the first auxiliary fuel valve can be manually increased to continue increasing the fuel flow in the primary combustion chamber, and then the temperature of the test item can be adjusted through the automatic control system of the first main fuel valve again to be stabilized at the corresponding temperature in the preset temperature curve.

[0072] As Figure 2 shown, the primary combustion chamber has air introduced into its air inlet end, and uses air and fuel for combustion to generate the first high-temperature gas that meets the first preset condition. The first high-temperature gas acts on the test turbine blade for a thermal shock test.

[0073] Specifically, when the requirements for the temperature adjustment range and temperature adjustment accuracy are relatively low, one combustion chamber can be selected as the combustion module. The gas pipeline passes air through the air inlet to promote the combustion of fuel in the primary combustion chamber, generating the first high-temperature gas with a relatively high temperature and outputting it to the test turbine blade. The preset condition can be the temperature range of the first high-temperature gas, which can vary according to different stages of the thermal shock test. For example: 200°C ± 30°C, 800°C ± 30°C, only as an example, but not limited thereto.

[0074] It should be noted that the combustion module can be provided with one combustion chamber or multiple combustion chambers arranged in sequence on the gas pipeline according to the actual situation. The more the number of combustion chambers, the larger the temperature range and the higher the accuracy of the finally output high-temperature gas. The control structure of each combustion chamber is the same as that of the primary combustion chamber in this embodiment (the fuel quantity is jointly controlled by the main fuel valve and the auxiliary fuel valve).

[0075] In some alternative embodiments, as Figure 3 shown, the fuel boosting module further includes: a second main fuel valve and a second auxiliary fuel valve. The combustion module further includes: a secondary combustion chamber.

[0076] As Figure 3 shown, for the second main fuel valve, its first end is connected to the outlet end of the fuel tank, and its second end is connected to the fuel inlet end of the secondary combustion chamber, and is used to change the opening degree according to the second instruction of the controller to adjust the fuel quantity entering the secondary combustion chamber. The second instruction is the opening degree change amount of the second main fuel valve obtained by the controller through data processing based on the real-time temperature and the preset temperature rising curve.

[0077] As Figure 3 shown, the second auxiliary fuel valve is connected in parallel with the second main fuel valve and is used to assist the second main fuel valve in controlling the fuel quantity entering the secondary combustion chamber.

[0078] As Figure 3 shown, for the secondary combustion chamber, its air inlet end is connected to the outlet end of the primary combustion chamber and the gas pipeline, and air and the first high-temperature gas are introduced. The air and fuel are used for combustion to heat the first high-temperature gas, generating the second high-temperature gas that meets the second preset condition. The second high-temperature gas acts on the test turbine blade for the thermal shock test.

[0079] Specifically, the second main fuel valve and the second auxiliary fuel valve are used to control the fuel quantity entering the secondary combustion chamber to change the real-time temperature of the second high-temperature gas output by the secondary combustion chamber, similar to using the first main fuel valve and the first auxiliary fuel valve to control the fuel quantity entering the primary combustion chamber to change the real-time temperature of the first high-temperature gas output by the primary combustion chamber in the previous embodiment.

[0080] When two - stage combustion chambers are sequentially connected to a gas pipeline, air enters the two - stage combustion chambers of the thermal shock test device through the air inlet of the gas pipeline respectively. The first high - temperature gas generated in the first - stage combustion chamber is introduced into the second - stage combustion chamber through the gas pipeline. At the same time, the second - stage combustion chamber also introduces new air through a branch of the air pipeline for auxiliary combustion to continuously increase the gas temperature. The second high - temperature gas generated by the combustion in the second - stage combustion chamber acts on the turbine blade for high - temperature thermal shock test. The fuel flow rate and fuel pressure entering the two - stage combustion chambers are controlled by the fuel boosting system, and the temperature of the tested turbine blade is controlled by changing the outlet temperature of the second - stage combustion chamber.

[0081] The temperature change range of the gas in the second - stage combustion chamber is relatively small. The second - stage combustion chamber is connected to the fuel boosting system through a valve. By controlling the valve opening, the fuel quantity entering the second - stage combustion chamber is adjusted, and then the temperature of the second high - temperature gas generated in the second - stage combustion chamber is accurately regulated. The second high - temperature gas is directly transmitted to the tested turbine blade, so the temperature of the second high - temperature gas should conform to the temperature change range of the actual service condition, that is, the second preset condition is the preset heating curve. Two - stage combustion chambers are designed to achieve the entire heating process. The first - stage combustion chamber makes a rough adjustment, and the second - stage combustion chamber makes a fine adjustment. High - temperature gas impact environment is provided by burning aviation kerosene in the two - stage combustion chambers. The intake section and exhaust section are designed to simulate the motion state of the high - temperature gas to ensure that the loading condition of the tested turbine blade during the test is basically the same as that in the engine during service.

[0082] In some alternative embodiments, as Figure 4 shown, the temperature acquisition module includes: a plurality of thermocouples, each thermocouple includes a plurality of temperature measurement points, and the plurality of temperature measurement points of the plurality of thermocouples are arranged in a concentric - circle distribution pattern on the gas outlet cross - section of the combustion module.

[0083] Specifically, combined with the structural characteristics of the gas outlet position of the combustion chamber, at the outlet of the combustion module, multiple thermocouples are used for temperature measurement. Each thermocouple is provided with a plurality of temperature measurement points, and the plurality of temperature measurement points of the plurality of thermocouples are arranged in a concentric - circle distribution pattern on the gas outlet cross - section of the combustion module. As Figure 4 shown, there are four thermocouples, and four evenly - distributed temperature measurement points on each thermocouple. A total of sixteen temperature measurement points are arranged in a concentric - circle distribution pattern. Using a four - layer concentric - circle temperature detection matrix with a total of sixteen temperature measurement points to detect the global temperature from the center to the edge of the combustion module outlet, and then the average value of the sixteen detected temperatures is calculated in real - time as the real - time temperature of the second high - temperature gas. The real - time temperature of the second high - temperature gas is equivalent to the real - time temperature of the tested turbine blade, and the real - time temperature is transmitted to the controller, which controls it to generate an instruction according to the relationship between the real - time temperature and the preset temperature curve, and adjusts the opening of the main fuel valve in real - time. The opening of the main fuel valve of the combustion module is continuously adjusted in the form of negative feedback, so as to improve the temperature control accuracy of the tested turbine blade.

[0084] The thermal shock test device for turbine blades under near-service conditions provided in this embodiment ensures the full combustion of aviation kerosene and reasonably regulates the temperature uniformity at the outlet of the secondary combustion chamber through separate commissioning of the single-stage combustion chamber and combined commissioning of the two-stage combustion chamber. By collecting the real-time temperatures at different positions at the outlet of the secondary combustion chamber and feeding them back to the controller for combined commissioning, the accuracy of temperature detection is improved. The two main fuel valves are mainly controlled automatically, and at the same time, auxiliary fuel valves are added to assist in regulating the temperatures of the two-stage combustion chamber. When the automatic control system is difficult to achieve the entire range of the set temperature curve, the auxiliary fuel valves are manually adjusted to increase the gas pressure in the two-stage combustion chamber, thereby raising the temperature upper limit of the combustion chamber and the test piece and expanding the allowable temperature range of the thermal shock test.

[0085] According to an embodiment of the present invention, there is provided an embodiment of a thermal shock test method for turbine blades under near-service conditions. It should be noted that the steps shown in the flowchart of the accompanying drawings can be executed in a computer system such as a set of computer-executable instructions. And although the logical order is shown in the flowchart, in some cases, the steps shown or described can be executed in a different order than here.

[0086] In this embodiment, a thermal shock test method for turbine blades under near-service conditions is provided, which can be used in the above computer system or the thermal shock test device for turbine blades under near-service conditions in the previous embodiment. Those that have been described will not be repeated. Figure 5 is a flowchart of the thermal shock test method for turbine blades under near-service conditions according to an embodiment of the present invention, as Figure 5 shown, the process includes the following steps:

[0087] Step S101, obtain the real-time temperature of the test turbine blade, and compare the real-time temperature with the set temperature corresponding to the preset temperature rise curve.

[0088] Specifically, the temperature acquisition module is used to collect the real-time temperature of the high-temperature gas at the outlet of the combustion module in real time as the real-time temperature of the test turbine blade. Since the high-temperature gas directly acts on the test turbine blade, the real-time temperature of the high-temperature gas can be used as the real-time temperature of the test turbine blade. To ensure that the temperature change process of the test turbine blade in the thermal shock test is the same as the actual working environment, it is necessary to control the change curve of the real-time temperature to change according to the preset temperature rise curve. Therefore, it is necessary to compare the real-time temperature with the set temperature corresponding to the preset temperature rise curve in order to control the fuel boosting module according to the comparison result.

[0089] The preset temperature rise curve is a temperature change curve of the turbine blade obtained through finite element calculation by combining with the actual engine test spectrum. According to the temperature change and holding time of the blade under different service conditions, a thermal shock test load spectrum is compiled. Based on the thermal shock test load spectrum, the temperature rise curve of the turbine blade is set, which can more accurately simulate the temperature load situation borne by the turbine blade during actual service.

[0090] Step S102, adjust the fuel quantity entering the combustion module according to the comparison result, change the real-time temperature of the tested turbine blade, so that the real-time temperature meets the preset conditions, and the temperature change curve of the tested turbine blade coincides with the preset temperature rise curve.

[0091] Specifically, according to the comparison result between the real-time temperature and the corresponding temperature in the preset temperature rise curve, adjust the fuel quantity entering the combustion module by controlling the fuel valve between the fuel boosting module and the combustion module, change the real-time temperature of the high-temperature gas at the outlet of the combustion module, so that the temperature change curve of the real-time temperature coincides with the preset temperature rise curve.

[0092] It should be noted that the combustion module can set a combustion chamber on the gas pipeline or set multiple combustion chambers in sequence according to the actual situation. The more the number of combustion chambers, the larger the temperature range and the higher the accuracy of the finally output high-temperature gas. The control structures of each combustion chamber are the same (the fuel quantity is jointly controlled by the main fuel valve and the auxiliary fuel valve). In this embodiment, a two-stage combustion chamber is taken as an example for illustration, but it is not limited thereto.

[0093] The thermal shock test method for turbine blades under near-service conditions provided in this embodiment ensures that the loading conditions of the tested turbine blades during the thermal shock test are basically the same as those inside the engine during the service state by designing the combustion of aviation kerosene in a two-stage combustion chamber, and uses the single-stage adjustment and two-stage joint adjustment of the two-stage combustion chamber to ensure the accuracy of the test temperature, so as to make the environment of the thermal shock test closer to the actual engineering situation and accurately evaluate the service life of the turbine blade and the coating at high temperatures.

[0094] In some optional implementation manners, the above step S101 includes:

[0095] Step S1011, obtain the real-time temperatures of multiple temperature measurement points on the cross section of the gas outlet of the combustion module.

[0096] Step S1012, calculate the real-time temperature of the high-temperature gas as the real-time temperature of the tested turbine blade according to the real-time temperatures of multiple temperature measurement points.

[0097] Specifically, to accurately collect the real temperature of the turbine blade under test, multiple temperature measurement points are set at the outlet of the secondary combustion chamber to detect the global temperature from the center to the edge of the cross-section at the outlet of the secondary combustion chamber. The real-time temperature of the second-highest temperature gas reaching the turbine blade under test can be calculated by taking the average value of multiple temperature measurement points in real time. Since the distance between the turbine blade under test and the outlet of the secondary combustion chamber is very small, the real-time temperature of the second-highest temperature gas can be used to represent the real-time temperature of the turbine blade under test.

[0098] Step S1013: Analyze the temperature fields of the turbine blades under multiple service conditions of the engine to obtain the preset temperature rise curve of the turbine blades.

[0099] Step S1014: Compare the real-time temperature of the turbine blade under test at the same sampling moment with the set temperature corresponding to it in the preset temperature rise curve to obtain the comparison result.

[0100] Specifically, by analyzing the temperature loads of the turbine blades under various service conditions such as engine idling, takeoff, and cruise, an engine test spectrum is obtained. As Figure 6 shown, it is a schematic diagram of the engine test spectrum, where the abscissa is time (min), the blue curve represents the change process of the turbine rotor speed (r / min) in different test stages, and the red curve represents the change process of the surface temperature (°C) of the turbine blade in different test stages. Then, the temperature field data under each service condition is determined through finite element analysis, and a thermal shock test load spectrum (equivalent to the preset temperature rise curve) is compiled according to the typical states of the engine. As Figure 7 shown, it is a schematic diagram of the preset temperature rise curve. The thermal shock test load spectrum can be used to determine the corresponding relationship between the temperature load state of the turbine blade and each service condition of the engine, the corresponding relationship between the holding time of each state and the duration of the engine state, the corresponding relationship between the temperature rise between each state and the time required for the engine state transformation, and the corresponding relationship between the temperature drop between each state and the time required for the engine state transformation. The specific analysis process can be realized by using relevant existing technologies and will not be elaborated here.

[0101] In some optional implementation manners, adjusting the fuel quantity entering the combustion module according to the comparison result in the above step S102 includes:

[0102] Step S1021: When the real-time temperature is less than the set temperature corresponding to it in the preset temperature rise curve, and the deviation between the real-time temperature and the set temperature corresponding to it in the preset temperature rise curve is greater than the upper limit of the preset deviation range, increase the fuel quantity entering the combustion module.

[0103] Step S1022: When the real-time temperature is greater than the set temperature corresponding to it in the preset temperature rise curve, and the deviation between the real-time temperature and the set temperature corresponding to it in the preset temperature rise curve is greater than the upper limit of the preset deviation range, reduce the fuel quantity entering the combustion module.

[0104] Specifically, the second high-temperature gas is heated up successively through combustion in two-stage combustion chambers. The temperature of the first high-temperature gas generated by combustion in the first-stage combustion chamber is basically close to the corresponding temperature in the preset temperature rise curve. For example, if the corresponding temperature in the preset temperature rise curve is 1000 °C, the temperature of the gas combusted in the first-stage combustion chamber can reach 800 °C - 900 °C. Then, the second-stage combustion chamber is used for fine regulation, and the second high-temperature gas with a temperature deviation within the preset deviation range from the corresponding temperature in the preset temperature rise curve needs to be generated through re-combustion in the second-stage combustion chamber. The preset deviation range can be [0 °C - 30 °C], which is only for example and not limited thereto. On the basis of ensuring good airtightness of the combustion chamber, the single-stage commissioning of the single-stage combustion chamber and the joint commissioning of the two-stage combustion chambers are carried out to ensure the full combustion of aviation kerosene and reasonably regulate the temperature uniformity of the second-stage combustion chamber outlet. Multiple temperature measuring points of multiple temperature measuring thermocouples are used to detect the temperature conditions at different positions at the outlet and feedback to the control system for joint commissioning. Through actual testing, when the temperature at the combustion chamber outlet reaches 1100 °C, the temperature distribution non-uniformity is less than 0.015.

[0105] In some alternative embodiments, increasing the fuel quantity entering the combustion module includes:

[0106] Increasing the opening degree of the main fuel valve in the fuel boosting module, and the fuel boosting module is used to supply fuel to the combustion module.

[0107] If the real-time temperature is less than the set temperature corresponding to the preset temperature rise curve when the opening degree of the main fuel valve is at its maximum, and the deviation between the real-time temperature and the set temperature corresponding to the preset temperature rise curve is greater than the upper limit of the preset deviation range, then increase the opening degree of the auxiliary fuel valve in the fuel boosting module until the real-time temperature meets the preset conditions.

[0108] Specifically, the process of adjusting the opening degrees of the main fuel valve and the auxiliary fuel valve in the two-stage combustion chambers has been described in detail in the previous embodiment and will not be elaborated here.

[0109] The thermal shock test method for turbine blades under near-service conditions provided in this embodiment starts from the actual engineering requirements, aims at the service environment state of turbine working blades, fully combines the actual service loads of the engine, obtains the temperature changes of turbine blades under different service states through finite element analysis, and compiles a test load spectrum equivalent to the actual service environment; designs the simulation of engine structural layouts such as two-stage combustion chambers, intake sections, and exhaust sections, provides a high-temperature gas environment for turbine blades to work by combusting aviation kerosene, and truly simulates the service environment characteristics of the blades; arranges multiple temperature measuring points evenly in a concentric circle layout, combines the advantages of the automatic control system for temperature commissioning, and ensures the accuracy and uniformity of the test temperature through single-stage regulation and two-stage joint regulation of the two-stage combustion chambers. Compared with the existing related technologies of thermal shock tests, this embodiment is closer to the actual engineering situation and is of great significance for accurately evaluating the service life of turbine blades and coatings at high temperatures.

[0110] An embodiment of the present invention further provides a computer device having the above-mentioned Figure 1 thermal shock test device for turbine blades under near-service conditions as shown.

[0111] Please refer to Figure 8 , Figure 8 which is a schematic structural diagram of a computer device provided by an alternative embodiment of the present invention. As Figure 8 shown, the computer device includes: one or more processors 10, a memory 20, and interfaces for connecting various components, including a high-speed interface and a low-speed interface. Each component communicates with each other using different buses and can be installed on a common motherboard or installed in other ways as needed. The processor can process instructions executed within the computer device, including instructions stored in the memory or on the memory to display graphical information of the GUI on an external input / output device (such as a display device coupled to the interface). In some alternative embodiments, if necessary, multiple processors and / or multiple buses can be used together with multiple memories and multiple memories. Similarly, multiple computer devices can be connected, and each device provides some necessary operations (for example, as a server array, a set of blade servers, or a multi-processor system). Figure 8 In

[0112] FIG., a single processor 10 is taken as an example.

[0113] The processor 10 can be a central processing unit, a network processor, or a combination thereof. Among them, the processor 10 can further include a hardware chip. The above-mentioned hardware chip can be an application-specific integrated circuit, a programmable logic device, or a combination thereof. The above-mentioned programmable logic device can be a complex programmable logic device, a field programmable gate array, a general array logic, or any combination thereof.

[0113] Wherein, the memory 20 stores instructions executable by at least one processor 10, so that the at least one processor 10 executes the method shown in the above embodiment.

[0114] The memory 20 can include a program storage area and a data storage area. Among them, the program storage area can store an operating system and application programs required for at least one function; the data storage area can store data created according to the use of the computer device. In addition, the memory 20 can include a high-speed random access memory and can also include a non-transitory memory, such as at least one magnetic disk storage device, a flash memory device, or other non-transitory solid-state storage devices. In some alternative embodiments, the memory 20 can optionally include a memory remotely set relative to the processor 10, and these remote memories can be connected to the computer device through a network. Examples of the above-mentioned network include but are not limited to the Internet, an enterprise intranet, a local area network, a mobile communication network, and combinations thereof.

[0115] The memory 20 may include volatile memory, such as random access memory; the memory may also include non-volatile memory, such as flash memory, a hard disk or a solid state drive; the memory 20 may further include a combination of the above types of memory.

[0116] The computer device further includes a communication interface 30 for the computer device to communicate with other devices or a communication network.

[0117] An embodiment of the present invention further provides a computer-readable storage medium. The method according to the embodiment of the present invention can be implemented in hardware, firmware, or be implemented as computer code that can be recorded on a storage medium, or be implemented by downloading through a network the original computer code stored in a remote storage medium or a non-transitory machine-readable storage medium and to be stored in a local storage medium, so that the method described herein can be stored in such software processing on a storage medium using a general-purpose computer, a dedicated processor, or programmable or dedicated hardware. Among them, the storage medium can be a magnetic disk, an optical disk, a read-only memory, a random access memory, a flash memory, a hard disk or a solid state drive, etc.; further, the storage medium may further include a combination of the above types of memory. It can be understood that a computer, a processor, a microprocessor controller or programmable hardware includes a storage component that can store or receive software or computer code, and when the software or computer code is accessed and executed by the computer, the processor or the hardware, the method shown in the above embodiment is implemented.

[0118] Although the embodiments of the present invention have been described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the present invention, and such modifications and variations all fall within the scope defined by the appended claims.

Claims

1. A thermal shock test device for turbine blades under near-service conditions, characterized in that: The device comprises: a fuel pressure boosting module, a controller, a combustion module, and a temperature acquisition module, wherein: A fuel boosting module, the outlet of which is connected to the fuel inlet of the combustion module, for providing fuel to the combustion module; The combustion module has air introduced into the air inlet end, and uses air and fuel to burn to generate high-temperature combustion gas that meets preset conditions. The high-temperature combustion gas acts on the turbine blades under test to perform a thermal shock test; A temperature acquisition module, which is used to acquire the real-time temperature of the high-temperature fuel gas at the outlet of the combustion module; The controller is used to obtain the real-time temperature of the high-temperature fuel gas, and control the fuel boost module according to the real-time temperature to change the amount of fuel entering the combustion module, so that the real-time temperature changes according to a preset temperature rise curve.

2. The thermal shock testing device according to claim 1, characterized in that: The fuel boost module includes: a fuel tank, a first main fuel valve, a first auxiliary fuel valve, and the combustion module includes: a primary combustion chamber, wherein: a fuel tank, the outlet ends of which are respectively connected to the first end of the first main fuel valve and the first end of the first auxiliary fuel valve, for providing fuel to the combustion module; a first main fuel valve, a second end of which is connected to the fuel inlet of the primary combustion chamber, and is used to change the opening according to a first instruction of the controller to adjust the amount of fuel entering the primary combustion chamber, wherein the first instruction is the opening change of the first main fuel valve obtained by the controller through data processing according to the real-time temperature and a preset temperature rise curve; a first auxiliary fuel valve, connected in parallel with the first main fuel valve, for assisting the first main fuel valve in controlling the amount of fuel entering the primary combustion chamber; The first combustion chamber has air introduced into the air inlet end, and utilizes the air and fuel to burn to generate the first high-temperature combustion gas that meets the first preset condition. The first high-temperature combustion gas acts on the turbine blade under test to perform a thermal shock test.

3. The thermal shock testing device according to claim 2, characterized in that: The fuel boost module further includes: a second main fuel valve, a second auxiliary fuel valve, and the combustion module further includes: a secondary combustion chamber, wherein: a second main fuel valve, a first end of which is connected to the outlet end of the fuel tank, and a second end of which is connected to the fuel inlet end of the secondary combustion chamber, and is used to change the opening according to a second instruction of the controller to adjust the amount of fuel entering the secondary combustion chamber, wherein the second instruction is the opening change of the second main fuel valve obtained by the controller through data processing according to the real-time temperature and a preset temperature rise curve; a second auxiliary fuel valve, connected in parallel with the second main fuel valve, for assisting the second main fuel valve in controlling the amount of fuel entering the secondary combustion chamber; The secondary combustion chamber has an air inlet end connected to the outlet end of the primary combustion chamber and the gas pipeline, and air and a first high-temperature fuel gas are introduced into the secondary combustion chamber. The air and fuel are used to burn and heat the first high-temperature fuel gas to generate a second high-temperature fuel gas that meets the second preset condition. The second high-temperature fuel gas acts on the turbine blade under test to perform a thermal shock test.

4. The thermal shock testing device according to claim 1, characterized in that: The temperature acquisition module comprises: a plurality of thermocouples, each of which comprises a plurality of temperature measuring points, and the plurality of temperature measuring points of the plurality of thermocouples are arranged in a concentric circle distribution manner on the gas outlet cross section of the combustion module.

5. A thermal shock test method for turbine blades under near-service conditions, characterized in that: The thermal shock test method is applied to the thermal shock test device described in any one of claims 1 to 4, and the method comprises: Obtaining the real-time temperature of the turbine blade under test, and comparing the real-time temperature with the corresponding set temperature in the preset heating curve; The amount of fuel entering the combustion module is adjusted according to the comparison result, and the real-time temperature of the tested turbine blade is changed so that the real-time temperature meets the preset conditions and the temperature change curve of the tested turbine blade is consistent with the preset temperature rise curve.

6. The thermal shock test method according to claim 5, characterized in that: The step of obtaining the real-time temperature of the turbine blade under test and comparing the real-time temperature with the corresponding set temperature in the preset heating curve includes: Obtaining the real-time temperature of multiple temperature measurement points on the gas outlet cross section of the combustion module; Calculating the real-time temperature of the high-temperature combustion gas according to the real-time temperatures of the plurality of temperature measuring points as the real-time temperature of the turbine blade under test; Analyze the temperature field of the turbine blades under multiple service states of the engine to obtain the preset temperature rise curve of the turbine blades; The real-time temperature of the tested turbine blade at the same sampling moment is compared with the corresponding set temperature in the preset temperature rise curve to obtain a comparison result.

7. The thermal shock test method according to claim 5, characterized in that: The step of adjusting the amount of fuel entering the combustion module according to the comparison result includes: When the real-time temperature is lower than the set temperature corresponding to the preset temperature rise curve, and the deviation between the real-time temperature and the set temperature corresponding to the preset temperature rise curve is greater than the upper limit of the preset deviation range, the amount of fuel entering the combustion module is increased; When the real-time temperature is greater than the set temperature corresponding to the preset temperature rise curve, and the deviation between the real-time temperature and the set temperature corresponding to the preset temperature rise curve is greater than the upper limit of the preset deviation range, the amount of fuel entering the combustion module is reduced.

8. The method according to claim 7, characterized in that The step of increasing the amount of fuel entering the combustion module comprises: increasing the opening of a main fuel valve in a fuel boost module, the fuel boost module being used to provide fuel to the combustion module; If the opening of the main fuel valve is at its maximum, the real-time temperature is lower than the corresponding set temperature in the preset temperature rise curve, and the deviation between the real-time temperature and the corresponding set temperature in the preset temperature rise curve is greater than the upper limit of the preset deviation range, then the opening of the auxiliary fuel valve in the fuel boosting module is increased until the real-time temperature meets the preset conditions.

9. A computer device, characterized in that: include: A memory and a processor, wherein the memory and the processor are communicatively connected to each other, the memory stores computer instructions, and the processor executes the method according to any one of claims 5 to 8 by executing the computer instructions.

10. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores computer instructions, and the computer instructions are used to enable a computer to execute the method according to any one of claims 5 to 8.

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