Method and device for determining pneumatic instability boundary of compression system in complete machine environment

Abstract

A method and device for determining a pneumatic instability boundary of a compression system in a complete machine environment, the method is applied to an engine test device, the engine test device comprises an engine body and a test modified turbine with a variable circulation sectional area, and the engine body comprises the compression system. The method comprises the steps that a compression system is controlled to enter a surge state by adjusting the circulation sectional area of a refitted turbine for testing and combining a fuel oil step method, or the compression system is controlled to enter the surge state by adjusting the circulation sectional area of the refitted turbine for testing and combining a high-pressure gas filling method; and determining an instability boundary point of the engine test device, and determining a pneumatic instability boundary of the compression system at the current rotating speed according to the instability boundary point. Therefore, the pneumatic instability boundary of the compression system in the whole machine environment is determined on the premise of not increasing the additional burden of a turbine and compression system components.

Description

Technical Field

[0001] This article relates to, but is not limited to, the field of engine technology, and in particular to a method and device for determining the aerodynamic instability boundary of a compression system in a complete machine environment. Background Art

[0002] Compression systems are widely used in the power and energy industries. While high-load operation improves performance, it also poses significant stability challenges. In particular, when operating near the instability boundary, surge is prone to occur, impacting performance and threatening safety. Accurately assessing this boundary is crucial for reducing development costs and improving reliability and performance. Because test results in a component environment may not be accurate for the entire system, full-system surge testing is typically required after prototype production to determine the true instability boundary.

[0003] In the related art, a fuel step method is usually used to obtain the instability boundary of the compression system in the whole machine environment.

[0004] However, although the fuel step method can effectively trigger compression system instability, the high temperature in the combustion chamber caused by it may pose a threat to the safety of the turbine structure. Summary of the Invention

[0005] The embodiments of the present application provide a method and apparatus for determining the aerodynamic instability boundary of a compression system in an overall machine environment, thereby achieving determination of the aerodynamic instability boundary of a compression system in an overall machine environment without adding additional burden to the turbine and compression system components.

[0006] The present invention provides a method for determining the aerodynamic instability boundary of a compression system in an entire engine environment, which is applied to an engine test device. The engine test device includes an engine body and a modified test turbine with a variable flow cross-sectional area. The engine body includes a compression system. The method includes: Controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a fuel step method, or controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a high-pressure gas charging method; An instability boundary point of the engine test device is determined, and an aerodynamic instability boundary of the compression system at a current rotational speed is determined based on the instability boundary point.

[0007] The embodiment of the present application further provides an engine test device, comprising: a memory and a processor; The memory is connected to the processor and is used to store programs; The processor is configured to implement the above-mentioned method for determining the aerodynamic instability boundary in the whole machine environment by running the program in the memory.

[0008] The embodiments of the present application include controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified test turbine in conjunction with a fuel ramp method, or by adjusting the flow cross-sectional area of ​​the modified test turbine in conjunction with a high-pressure gas injection method; determining the instability boundary point of the engine test device, and determining the aerodynamic instability boundary of the compression system at the current speed based on the instability boundary point. Thus, the aerodynamic instability boundary of the compression system is determined in a complete engine environment without adding additional burden to the turbine and compression system components.

[0009] Other features and advantages of the present application will be described in the following description, and in part will become apparent from the description, or will be understood by practicing the present application. Other advantages of the present application can be realized and obtained by the solutions described in the description and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] The accompanying drawings are used to provide an understanding of the technical solution of the present application and constitute a part of the specification. Together with the embodiments of the present application, they are used to explain the technical solution of the present application and do not constitute a limitation on the technical solution of the present application.

[0011] Figure 1 This is a flow chart of a method for determining the aerodynamic instability boundary of a compression system in a whole-machine environment according to an embodiment of the present application; Figure 2 This is a flow chart of controlling a compression system to enter a surge state according to an embodiment of the present application; Figure 3 This is a flow chart of forcing a compression system to surge until the compression system enters a surge state according to an embodiment of the present application; Figure 4 This is a schematic diagram of the process of oil step coefficient adjustment operation according to an embodiment of the present application; Figure 5 This is another flow chart of controlling a compression system to enter a surge state according to an embodiment of the present application; Figure 6 This is another flow chart of forcing a compression system to surge until the compression system enters a surge state according to an embodiment of the present application; Figure 7 A schematic structural diagram of an engine test device provided in an embodiment of the present application; Figure 8 This is a schematic diagram of flow rate changes during a surge process according to an embodiment of the present application; Figure 9 Schematic diagram of a slope change process of a flow rate change curve during a surge process according to an embodiment; Figure 10 A schematic diagram of the surge dynamic process according to an embodiment of the present application; Figure 11 This is a flow chart of another method for determining the aerodynamic instability boundary of a compression system in a whole machine environment according to an embodiment of the present application; Figure 12 This is a flow chart of another method for determining the aerodynamic instability boundary of a compression system in a whole machine environment according to an embodiment of the present application; Figure 13 This is a structural schematic diagram of an engine testing device according to an embodiment of the present application. DETAILED DESCRIPTION

[0012] This application describes multiple embodiments, but this description is exemplary rather than restrictive, and it is obvious to those skilled in the art that there may be more embodiments and implementations within the scope of the embodiments described in this application. Although many possible feature combinations are shown in the drawings and discussed in the detailed description, many other combinations of the disclosed features are also possible. Unless specifically limited, any feature or element of any embodiment may be used in combination with any other feature or element in any other embodiment, or may replace any other feature or element in any other embodiment.

[0013] The present application includes and contemplates combinations of features and elements known to those of ordinary skill in the art. The embodiments, features, and elements disclosed in this application may also be combined with any conventional features or elements to form a unique inventive solution. Any features or elements of any embodiment may also be combined with features or elements from other inventive solutions to form another unique inventive solution. Therefore, it should be understood that any feature shown and / or discussed in this application may be implemented individually or in any appropriate combination. Therefore, except for the limitations made according to the appended claims and their equivalents, the embodiments are not subject to other limitations. In addition, various modifications and changes may be made within the scope of protection of the appended claims.

[0014] In addition, when describing representative embodiments, the specification may have presented the method and / or process as a specific sequence of steps. However, to the extent that the method or process does not rely on the specific order of the steps described herein, the method or process should not be limited to the steps in the specific order described. As will be understood by those skilled in the art, other orders of steps are also possible. Therefore, the specific order of the steps set forth in the specification should not be interpreted as a limitation to the claims. In addition, the claims for the method and / or process should not be limited to performing their steps in the order written, and those skilled in the art can readily understand that these orders can be changed and still remain within the spirit and scope of the embodiments of the present application.

[0015] Rotating compression components, represented by compressors, are widely used in the aviation and energy industries. Due to their own structure and performance requirements, compressors usually need to operate under high load conditions to achieve higher working efficiency. However, under high-load conditions, the compressor is very likely to enter a state of aerodynamic instability. Once instability occurs, the compressor will vibrate violently, its working performance will drop sharply, and it may cause serious structural damage. The scope of impact may extend to the entire power system and even the engine itself, thereby endangering the safety of the driven equipment and other related systems. Therefore, accurately obtaining the aerodynamic instability boundary of the compressor has always been a research direction that the industry focuses on. Reducing the technical and economic costs of this process will significantly improve the R&D efficiency and engineering application level of related industries.

[0016] The aerodynamic instability boundary of a compressor differs significantly between an ideal component environment and an actual operating environment. During engine development, in order to obtain the aerodynamic instability boundary of the engine compressor as close to the actual operating environment as possible, it is often necessary to conduct a whole-machine gas pressure test on the engine in a whole-machine environment. Whole-machine gas pressure tests usually use a fuel step method or a high-pressure gas filling method to obtain the aerodynamic instability boundary of the compressor. Among them, the fuel step method will greatly increase the temperature before the engine turbine, significantly worsening the turbine operating environment, thereby significantly reducing the turbine life and even causing turbine damage. The high-pressure gas filling method requires a stable air source support, which often requires a major modification of the engine whole-machine test bench. Both methods are costly to use.

[0017] To this end, an embodiment of the present disclosure provides a method for determining an aerodynamic instability boundary in a complete machine environment, which is applied to an engine test device. The engine test device includes: an engine body and a modified test turbine with a variable flow cross-sectional area. The engine body includes a compression system, such as Figure 1 As shown, the method includes: Step 100: Controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a fuel step method, or by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a high-pressure gas charging method; Step 110: Determine the instability boundary point of the engine test device, and determine the aerodynamic instability boundary of the compression system at the current speed based on the instability boundary point.

[0018] The modified test turbine with a variable flow cross-sectional area in the embodiment of the present application can be obtained by any suitable means. One implementation method can refer to the test turbine structure described in patent application number 2025103302586. This patent discloses the relevant design and adjustment mechanism in detail, which can be used as a reference and applied to the modified test turbine in this embodiment. In this patent, the stator blades of the test turbine are configured as rotatable blades to reduce the flow cross-sectional area of ​​the test turbine by rotating the stator blades. The test turbine also includes an actuating device connected to the stator blades of the test turbine and configured to drive the stator blades to rotate. The actuating device includes: a plurality of rotating shafts and a driving mechanism; the number of stator blades is multiple, and the plurality of rotating shafts are arranged one-to-one with the plurality of stator blades of the test turbine; each rotating shaft extends in a radial direction of the test turbine and is connected to a corresponding stator blade, configured to drive the corresponding stator blade to rotate about the rotating shaft; the driving mechanism is connected to the plurality of rotating shafts and configured to drive the plurality of rotating shafts to rotate synchronously to synchronously adjust the working angles of the plurality of stator blades.

[0019] The method for determining the aerodynamic instability boundary in a whole-machine environment provided in an embodiment of the present application replaces the original turbine with a modified turbine for testing and installs it in the corresponding position of the engine body, while the structure of the engine body is the original structure of the engine.

[0020] The present invention provides a method for determining the aerodynamic instability boundary in a complete engine environment. This method controls the compression system into a surge state by adjusting the flow cross-sectional area of ​​a modified test turbine and combining it with a fuel step method, or by adjusting the flow cross-sectional area of ​​the modified test turbine and combining it with a high-pressure gas injection method. The method also determines the instability boundary point of the engine test device and, based on the instability boundary point, determines the aerodynamic instability boundary of the compression system at the current speed. This method thus achieves the determination of the aerodynamic instability boundary of the compression system in a complete engine environment without adding additional burden to the turbine and compression system components.

[0021] In related technologies, a single fuel step surge method is usually used to achieve compression system instability. However, this method causes the combustion chamber temperature to rise sharply, worsening the working environment of the turbine, and may further cause thermal damage to the engine structure, posing a significant safety risk. In contrast, the method for determining the aerodynamic instability boundary in a whole-machine environment provided in the embodiment of the present application does not rely solely on the fuel step surge method. Instead, it controls the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine used for the test and combining the fuel step method. Therefore, there is no need to inject a large amount of fuel into the combustion chamber to significantly increase its temperature and pressure. Only a small amount of fuel needs to be injected to trigger compression system instability, thereby effectively reducing the thermal load and structural damage risk during the test and significantly improving the safety and reliability of the test.

[0022] In related technologies, high-pressure gas injection is often used to induce compression system instability, but this method requires a large amount of high-pressure gas source, which is expensive to obtain. In contrast, the method for determining the aerodynamic instability boundary in a complete machine environment provided in the embodiments of the present application controls the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified test turbine and combining it with the high-pressure gas injection method. Therefore, there is no need to inject a large amount of high-pressure gas source into the combustion chamber to significantly increase its pressure. Only a small amount of high-pressure gas source is required to trigger compression system instability, thereby effectively reducing the demand for high-pressure gas source and, in turn, its acquisition cost.

[0023] The compression system includes a compressor and, if necessary, may also include a fan and an air inlet on the upstream side of the compressor and a matching piping system.

[0024] In an exemplary embodiment, when the compression system is controlled to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining the fuel step method, as shown in FIG. Figure 2 As shown in the figure, by adjusting the flow cross-sectional area of ​​the modified turbine used in the test and combining it with the fuel step method, the compression system is controlled to enter a surge state, including: Step 200: Rotate the stator blades of the modified turbine for testing to a calibrated initial position; at the initial position, the stator blades of the modified turbine for testing are in the same position as the stator blades of the unmodified turbine; Step 210: Ignite and start the engine test device, and accelerate it to a set speed for the whole-machine breathing test; Step 220: Control the stator blades of the modified turbine for the test to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test, and combine the fuel step method to force the compression system to surge until the compression system enters a surge state.

[0025] During the implementation of step 200, the stator 0° position of the modified turbine (also called a variable geometry turbine) used for the test may be calibrated according to the original geometry of the engine so that the angle of the variable geometry turbine is equivalent to the original geometry. The position of the actuator ring is repeatedly changed to calibrate the correspondence between the position of the actuator and the angle of the variable geometry turbine blades, and the calibrated variable geometry turbine angle is set to 0°.

[0026] During the implementation of step 210, the engine can be ignited and started while ensuring that the variable geometry turbine is adjusted to 0°. The engine starting control law before modification is used to ignite the engine and push it to the idle speed. The variable geometry turbine angle is kept unchanged, and the engine acceleration control law before modification is used to gradually accelerate the engine until the engine reaches the speed required to be measured in this test.

[0027] During the implementation of step 220, the engine can be maintained at a speed required for the surge test, and the engine variable geometry turbine angle can be gradually reduced in steps of 1°. During the process of reducing the variable geometry angle, the operating state of the engine should be kept roughly unchanged. That is, each time the variable geometry angle is fully adjusted, a certain amount of time should be waited until the typical performance parameters of the engine, such as the speed and the total pressure ratio, have not changed significantly before making the next adjustment.

[0028] In an exemplary embodiment, Figure 3 As shown, the stator blades of the modified turbine for the test are controlled to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test, and the compression system is forced to surge by combining the fuel step method until the compression system enters a surge state, including: Step 300: Continuously adjust the fuel step coefficient b To make the fuel flow rate follow b The ratio of times changes step by step, according to the final determined fuel step coefficient b Calculate the flow cross-sectional area required by the modified turbine for the test, and control the stator blades of the modified turbine for the test to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test to the calculated flow cross-sectional area; wherein, the fuel step coefficient adjusted at any time b Less than the fuel step coefficient a , a The fuel step coefficient that can cause the compression system to enter a surge state by using the fuel step method alone; Step 310: When the flow cross-sectional area of ​​the modified turbine for the test reaches the calculated flow cross-sectional area, use the final determined fuel step coefficient b After forcing the system to surge, determine whether the compression system has entered a surge state; Step 320: In response to the judgment result that the compression system has not entered the surge state, the stator blades of the modified turbine for the test are controlled to rotate gradually to continue to reduce the flow cross-sectional area, and after each reduction in the flow cross-sectional area, the fuel step coefficient finally determined is used. b Perform a fuel step and determine whether the compression system enters a surge state until the compression system enters a surge state.

[0029] For the engine before and after variable geometry adjustment, assuming that the engine before adjustment uses a fuel step to force the engine into surge, the fuel step coefficient (i.e. the ratio of the engine fuel injection amount before and after the fuel step) that can make the engine enter the surge state is a , the fuel step coefficient after the geometry change is b The fuel step coefficient and the flow cross-sectional area (also known as the turbine throat area) can be approximately approximated as: Where, is the turbine throat area before changing the geometry, is the turbine throat area after variable geometry, and the relationship between the change in the variable geometry turbine throat and the variable geometry angle is usually calibrated specifically for different cases through simulation or experimental methods, but a relatively simple approximate formula can also be used, such as: in, is the change in turbine throat area; is the chord length of the turbine stator blade; is the number of turbine stator blades; is the relative angle through which the turbine stator blades rotate; is the radius of the centerline of the turbine stator blade flow channel.

[0030] In an exemplary embodiment, the fuel step coefficient is continuously adjusted b To make the fuel flow rate follow b The ratio of the times undergoes a step change, including: Set the initial fuel step coefficient for fuel flow b and set the initial fuel step coefficient b As the current fuel step coefficient b , perform the following fuel step coefficient adjustment operation until the fuel flow rate is in accordance with b A step change in the ratio of times: The oil step coefficient adjustment operation is as follows Figure 4 Shown, including: Step 400: According to the current fuel step coefficient b Adjust the flow cross-sectional area of ​​the modified turbine used for the test and increase the amount of fuel supplied to the combustion chamber; Step 410: Determine whether the oil supply system can be adjusted according to the flow cross-sectional area. b The fuel flow rate can be changed in steps by a factor of 2. Step 420: If you are unable to b When the fuel flow rate step change is achieved at a ratio of times, continue to adjust the fuel step coefficient b and adjust the fuel step coefficient b As updated current fuel step factor b , continue with the fuel step coefficient adjustment operation.

[0031] When the fuel step is forced, the angle of the variable geometry turbine guide vane should be kept unchanged, and the fuel flow rate should be reduced from the initial normal working state to the fuel flow rate in a short time through the engine control system. Adjust to The fuel flow rate change time, here refers to the difference between the time when the actual fuel nozzle flow rate reaches the target flow rate and the time when the fuel nozzle flow rate starts to rise, rather than the time taken for the control command to be issued, should not exceed 1s. If within 1s, the engine is restricted by the engine control system or the fuel supply system, the engine cannot achieve the required b If the fuel step is equal to times the fuel flow rate, adjust the fuel step coefficient b .

[0032] In an exemplary embodiment, when the compression system is controlled to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a high-pressure gas charging method, the engine test device further includes: an air storage chamber, a heating device, the compression system includes: a compressor, and the engine body includes: a combustion chamber; By adjusting the flow cross-sectional area of ​​the modified turbine used in the test and combining the high-pressure gas filling method, the compression system is controlled to enter a surge state, such as Figure 5 Shown, including: Step 500: Rotate the stator blades of the modified turbine for testing to a calibrated initial position; at the initial position, the stator blades of the modified turbine for testing are in the same position as the stator blades of the unmodified turbine; Step 510: Ignite and start the engine test device, and accelerate it to a set speed for the whole-machine breathing test; Step 520: Control the stator blades of the modified test turbine to rotate gradually to reduce the flow cross-sectional area of ​​the modified test turbine, and combine the high-pressure gas charging method to force the compression system to surge until the compression system enters a surge state; wherein, the high-pressure gas charging method is a method of using a heating device to heat the high-pressure gas from the compressor in the air storage chamber and output the high-heat and high-pressure gas to the combustion chamber.

[0033] In an exemplary embodiment, a plurality of vent holes are provided on the outer wall of the flow channel at the compressor outlet, and the air storage cavity is connected to the vent holes by a vent pipe, and a valve is provided on the vent pipe; Control the stator blades of the modified test turbine to rotate gradually to reduce the flow cross-sectional area of ​​the modified test turbine, and combine the high-pressure gas charging method to force the compression system to surge until the compression system enters a surge state, including: Set the initial flow cross-sectional area required for the modified turbine used in the test, set the initial heating temperature for the heating device to heat the air storage chamber, use the set initial flow cross-sectional area as the current flow cross-sectional area, use the set initial heating temperature as the current heating temperature, and perform the following surge operation until the compression system enters a surge state; Breathing operation Figure 6 Shown include: Step 600: close the valve and use the heating device to heat the high-pressure gas in the gas storage chamber at the current heating temperature to generate high-heat high-pressure gas; Step 610: Open the valve to allow the high-temperature and high-pressure gas in the gas storage chamber to flow into the combustion chamber, thereby forcing the compression system to surge. During the surge process, determine whether the compression system enters a surge state. Step 620: In response to the judgment result that the compression system has not entered the surge state, reduce the current flow cross-sectional area to obtain an updated current flow cross-sectional area and / or increase the current heating temperature to obtain an updated current heating temperature, and continue to perform the surge operation using the updated current flow cross-sectional area and / or the updated current heating temperature.

[0034] In the implementation of step 620, in response to the judgment result that the compression system has not entered the surge state, the current flow cross-sectional area can be reduced to obtain an updated current flow cross-sectional area, and the surge operation can be performed using the updated current flow cross-sectional area; or the current heating temperature can be increased to obtain an updated current heating temperature, and the surge operation can be continued using the updated current heating temperature; or the current flow cross-sectional area can be reduced to obtain an updated current flow cross-sectional area, and the current heating temperature can be increased to obtain an updated current heating temperature, and the surge operation can be continued using the current flow cross-sectional area and the updated current heating temperature.

[0035] In related technologies, when using a high-pressure gas source method, it is usually necessary to make large-scale modifications to the tester and its peripheral equipment, such as adding a compressor and other devices to provide a gas source that meets the requirements. Such modifications not only involve high costs, but also require a long implementation period, and the overall feasibility is low. In contrast, the method for determining the aerodynamic instability boundary under the whole machine environment provided in the embodiment of the present application is to provide a plurality of vents on the outer wall of the flow channel at the outlet of the compressor, and the air storage chamber is connected to the vents (also called intake / exhaust interfaces) through a vent pipe. A valve is provided on the vent pipe to connect the air storage chamber to the vent through the pipe. The engine test device provided in the embodiment of the present application can be as follows: Figure 7 As shown, it includes: an air storage chamber 11, a heating device 12, an air vent 13 (only one is used as an example in the figure), a modified test turbine 14 with a variable flow cross-sectional area, and an electronically controlled valve 15. After the high-pressure gas generated by the compressor enters the air storage chamber, the heating device is used to heat the gas in the air storage chamber to generate high-temperature and high-pressure gas. When the valve is opened, the high-temperature and high-pressure gas enters the combustion chamber through the air vent, thereby increasing the outlet back pressure of the compression system and achieving pressure control of the compression system. The method of the embodiment of the present application does not require large-scale modification of existing equipment, and can be completed only through local structural optimization. It has the advantages of convenient modification, low cost and short implementation period.

[0036] In the embodiment of the present application, the valve provided on the ventilation line is required to be able to be controlled to open during the operation of the engine, and the actuation frequency of the valve should be no less than 10 Hz.

[0037] The material and structure of the air storage chamber must be able to withstand high temperatures and high pressures. The ventilation line needs to be able to accommodate both forward and reverse flow. When storing air, the stored gas can be heated using heating equipment while the valve is closed. However, since the airflow at the outlet of the engine compressor often already has a relatively high initial temperature, attention should be paid to the heating efficiency when selecting the heating method. For engines with a lower total pressure ratio, conventional resistance heating can be used. For high-design horizontal compressors with a higher pressure ratio, high-performance heating methods such as arc heating are usually required.

[0038] In one exemplary embodiment, the minimum area of ​​a single vent hole is as follows: in, is the minimum area of ​​a single vent, is the outlet mass flow rate when the compressor is working stably at the test working point, is the static temperature at the outlet of the compressor when it is working stably at the test working point, is the static pressure at the outlet of the compressor when it is working stably at the test working point, is the correction coefficient, which is determined according to the selection of valve and vent. is the number of vents, is the specific heat ratio, is the gas constant, is the temperature of the gas in the gas storage chamber.

[0039] For the vent holes, their size and number should be such that at the moment the valve is opened, the gas mass flow rate through the vent holes can meet the mass flow rate required for the test engine to breathe in. Generally speaking, this mass flow rate is 1 / 3 of the compressor outlet mass flow rate when the engine operating point is stable at the test speed.

[0040] At the same time, considering that the vent hole needs to maintain a certain flow rate for a certain period of time to allow the engine to surge, after determining the vent hole area, it is still necessary to ensure that there is enough air in the air storage chamber to ensure that the device can maintain the compressor at the surge boundary for a certain period of time, so that the compressor can surge smoothly. Since the typical surge characteristic time is above 10Hz, an example is provided here. To maintain a certain flow rate within 1s, the following should be met: in, is the gas constant, is the volume of the air storage chamber.

[0041] This formula requires a numerical method to solve for the volume of the air storage chamber. To minimize the burden of heating the gas in the air storage chamber, the volume of the air storage chamber should be as large as possible, and it is recommended to be at least 10 times the equivalent volume of the engine combustion chamber.

[0042] In an exemplary embodiment, the heating temperature of the heating device is as follows: and ; ; in, is the temperature of the gas storage chamber after heating, is the combustion chamber inlet gas temperature, is the combustion chamber inlet pressure, is the inlet pressure of the combustion chamber when it reaches the aerodynamic instability boundary during the gasping process. is the equivalent aerodynamic volume of the combustion chamber, is the aerodynamic volume of the air storage chamber, is the calorific value of fuel, is the specific heat capacity of the gas, is a pre-set fuel step coefficient, and Less than a The fuel step coefficient, a is a fuel step coefficient capable of causing the compression system to enter a surge state by using a fuel step method alone; is the air flow rate at the compressor outlet, It is the fuel flow rate when the engine is in stable working condition.

[0043] In an exemplary embodiment, whether the engine enters the surge state is determined based on whether the operating information of the engine test device satisfies the determination condition for entering the surge state; The operation information of the engine test device includes: physical quantities used to characterize the performance of the compression system, and signals used to characterize the operating status of the engine test device; The judgment conditions for entering the surge state include at least one of the following: a physical quantity used to characterize the performance of the compression system meets a set surge condition, and a signal used to characterize the operating state of the engine test device shows a set surge signal.

[0044] The setting of the gasping condition includes at least one of the following: the curvature change of the physical quantity satisfies a first set instability boundary condition, and the frequency of the periodic signal extracted according to the physical quantity satisfies a second set instability boundary condition; The gasping signal is configured to include at least one of a light signal and an acoustic signal. The acoustic signal may include, but is not limited to, a low-frequency humming sound and / or a continuous popping sound that is different from normal operation. The optical signal may include, but is not limited to, a flame emanating from the outlet of the engine test device.

[0045] In some exemplary embodiments, data corresponding to a physical quantity (such as the outlet pressure of the compressor) is selected from the data corresponding to the physical quantity used to characterize the performance of the compression system, and the physical quantity is determined. P A curve that changes with time. Figure 8 As shown in the figure, taking flow as an example, the change process of the physical quantity during the surge process is shown, which is characterized by periodic fluctuation accompanied by alternating positive and negative fluctuations. The curvature calculation formula is used to calculate the curve at each moment. t The curvature at Figure 9 shown.

[0046] For example: The average static pressure of the outlet section of the selected compressor is P The average static pressure is P , can be obtained by the following method: averaging the dynamic pressure signals collected by multiple outlet pressure probes in a circumferential direction (i.e., summing and averaging), and filtering them with a filter function (filtering out high-frequency and low-frequency signals). On this basis, the average static pressure can be obtained as P The curve changes with time, and the curvature calculation formula can be used to calculate the curve at each moment t The curvature of the .

[0047] The curvature calculation formula is: ; In the above formula, Representing physical quantities P The slope of the change curve, that is, the physical quantity P the speed of change; W Indicates the window length for calculating the slope, which is a positive integer; Representing physical quantities P exist W The amount of change within a certain period of time; f Indicates the sampling frequency.

[0048] The first set of instability boundary conditions includes: ; In the above formula, Indicates the time N Time physical quantity P The slope of A , A represents the threshold value of the compressor 11 with different configurations, Indicates the flow rate of the compressor. The physical meaning of this formula is: when the physical quantity P The speed of the change process reaches A , and the compressor flow rate is positive at this time, it is considered that the instability boundary is reached, and this operating condition is taken as the instability boundary point of the compressor.A The value depends on the compressor configuration of the engine being calculated (e.g., axial-flow, centrifugal, or combined axial-flow, centrifugal), as well as the engine size and power rating. In practice, the value is typically selected based on existing examples of engines with similar configurations, sizes, and power ratings.

[0049] In some exemplary embodiments, the frequency of the periodic signal extracted from the physical quantity is recorded as . The extraction method is as follows: using real-time fast Fourier transform, the main periodic signal frequencies in the selected physical quantity are extracted.

[0050] For example: The average static pressure of the outlet section of the selected compressor is P The average static pressure is P , can be obtained by the following method: averaging the dynamic pressure signals collected by multiple outlet pressure probes in a circumferential direction (i.e., summing and averaging), and filtering them with a filter function (filtering out high-frequency and low-frequency signals). Using real-time fast Fourier transform, the main periodic signal frequency in the dynamic pressure signal is extracted, which is .

[0051] The second set of instability boundary conditions includes: ; In the above formula, is a constant ， It is usually selected based on the specific configuration, size and power level of the engine compressor. The Helmholtz frequency of the engine test device under the current state is predicted and calculated. The meaning of this formula is that when the main periodic frequency of the dynamic change of the compressor outlet pressure is close to the Helmholtz frequency , it means that “the frequency of the periodic signal extracted according to the physical quantity satisfies the second set instability boundary condition” is established.

[0052] The calculation formula is: In the above formula, represents the speed of sound at the compressor outlet, is the equivalent cross-sectional area of ​​the compressor outlet, is the equivalent volume of the combustion chamber, is the equivalent length of the combustion chamber.

[0053] Figure 10The figure shows the instability boundary of the whole machine obtained by the method of the embodiment of the present application. The black track in the figure is the entire process of the surge condition of the compressor under the whole machine environment obtained by the method of the embodiment of the present application, and the gray circle is the instability boundary point determined according to this method.

[0054] The present application also provides a method for determining the aerodynamic instability boundary in a complete engine environment. The method is a composite method of a variable geometry turbine and a fuel step: using a variable geometry turbine to improve the fuel step effect, thereby reducing the requirement for combustion chamber temperature rise during the fuel step process, thereby reducing the damage to the engine caused by the fuel step. Figure 11 As shown, the specific process includes: Step 700: Install the engine on the test bench and set it to the initial position.

[0055] Install the modified engine on the test bench and secure the external air reservoir in place. Calibrate the initial position of the variable geometry turbine. Based on the original engine geometry, calibrate the modified variable geometry turbine stator to 0° so that the variable geometry turbine angle matches the original geometry. Repeatedly adjust the position of the actuator ring to calibrate the relationship between the actuator position and the variable geometry turbine blade angle. Set the calibrated variable geometry turbine angle to 0°.

[0056] Step 710: Set the engine to a test speed.

[0057] Ensure the variable turbine geometry is set to 0° and start the engine. Use the pre-modified engine starting control scheme to ignite the engine and bring it to idle speed. Maintaining the variable turbine geometry, gradually accelerate the engine using the pre-modified engine acceleration control scheme until the engine reaches the desired speed for this test.

[0058] Step 720: Adjust the variable geometry turbine angle to reduce the engine throat area.

[0059] Keep the engine at the speed required for surge testing, and gradually reduce the variable geometry turbine angle of the engine with a variable geometry angle of 1° as a step. In the process of reducing the variable geometry angle, the working state of the engine should be kept roughly unchanged, that is, each time the variable geometry angle is adjusted, wait for a period of time until the typical performance parameters of the engine, such as speed and total pressure ratio, have not changed significantly before making the next adjustment. For the engine before and after the variable geometry angle adjustment, assuming that the engine before the adjustment uses a fuel step to force the engine into a surge state, the fuel step coefficient (that is, the ratio of the engine fuel injection amount before and after the fuel step) that can cause the engine to enter a surge state is a , the fuel step coefficient after the geometry change is b The fuel step coefficient and turbine throat area can be approximately approximated as: Where, is the turbine throat area before changing the geometry, is the turbine throat area after variable geometry. The relationship between the change in the variable geometry turbine throat and the variable geometry angle usually requires a lot of actual calculations and calibration, but a relatively simple approximate formula can also be used, such as: Where, is the change in turbine throat area; is the chord length of the turbine stator blade; is the number of turbine stator blades; is the relative angle through which the turbine stator blades rotate; is the radius of the centerline of the turbine stator blade flow channel.

[0060] Step 730: Dynamically collect compressor outlet pressure parameters as surge judgment criteria.

[0061] like Figure 7 As shown, a dynamic pressure sensor 16 (also called a pressure probe) can be placed at the engine compressor outlet to determine whether the compressor has entered a surge state. The compressor outlet pressure parameters can be collected using the dynamic pressure sensor. Considering that the engine is in normal operating condition during the test, high-temperature combustion gas in the combustion chamber may flow back to the engine end during the surge process, so special consideration must be given to the selection of the dynamic pressure sensor. The surge frequency is approximately 100 Hz, and the frequencies of the pressure probe and the acquisition device are 20 kHz and 200 kHz, respectively, which meet the requirements for measuring surge phenomena. The dynamic pressure probe's measurement and acquisition system is started synchronously with engine startup. The results of the dynamic pressure acquisition system are monitored at all times.

[0062] Step 740: Fuel step pressure After the variable geometry angle adjustment is completed, the engine is subjected to fuel step pressure. When the fuel step pressure is applied, the angle of the variable geometry turbine guide vanes should be kept unchanged, and the fuel flow rate is reduced from the initial normal working state to the fuel flow rate in a short time through the engine control system. Adjust to The fuel flow rate change time, here refers to the difference between the time when the actual fuel nozzle flow rate reaches the target flow rate and the time when the fuel nozzle flow rate starts to rise, rather than the time taken for the control command to be issued, should not exceed 1s. If within 1s, the engine is restricted by the engine control system or the fuel supply system, the engine cannot achieve the required b If the fuel step is equal to the fuel flow rate, the process should return to step 720 and adjust the target fuel step coefficient. b , and adjust the variable geometry turbine throat area according to the formula Step 750: Determine whether the engine is breathing heavily based on the following conditions.

[0063] If at least one of P1, P2, and P3 is satisfied, the engine is determined to have entered a surge state, and the process proceeds to step 760. If none of P1, P2, and P3 is satisfied, the process returns to step 720, further reducing the variable-geometry turbine angle and then performing a fuel step surge. This cycle continues until the engine enters a surge state.

[0064] P1 conditions: The dynamic pressure signal measured at the compressor outlet is averaged circumferentially, and the high-frequency and low-frequency signals are filtered out using a filter function. The processed dynamic pressure signal is recorded as P Calculate on this basis P The curvature of the change process The process of change, in which: Where, Representing physical quantities P The slope of the change curve, that is, the physical quantity P the speed of change; W Indicates the window length for calculating the slope, which is a positive integer; Representing physical quantities P exist W The amount of change within a certain period of time; f Indicates the sampling frequency.

[0065] Determine whether the stability boundary conditions are met. The judgment conditions are: Where, Indicates the time N Time physical quantity P The slope of A ,in A represents the threshold value; Represents the compressor flow rate. The physical meaning of the above formula is that when the physical quantity P The speed of the change process reaches A When , and the compressor flow is positive at this time, it is considered that the surge boundary is reached, and this operating condition is taken as the surge boundary point of the compressor. A The value depends on the compressor configuration of the aircraft engine being calculated (e.g., axial-flow compressor, centrifugal compressor, combined axial-flow centrifugal compressor), as well as the engine size and power rating. In practice, the value is typically selected based on existing examples of engines with similar configurations, sizes, and power ratings.

[0066] P2 conditions: The dynamic pressure signal measured at the compressor outlet is averaged circumferentially, and the high-frequency and low-frequency signals are filtered out using a filter function. The main periodic signal frequencies in the dynamic pressure signal are extracted using a real-time fast Fourier transform. .

[0067] Based on the current working state of the engine, the engine flow rate, pressure and other parameters obtained through measurement or model calculation are combined with the engine's own geometric characteristic parameters to calculate and predict the Helmholtz frequency of the engine entering the surge state at this time. . The calculation method is: in represents the speed of sound at the compressor outlet, is the equivalent cross-sectional area of ​​the compressor outlet, is the equivalent volume of the combustion chamber, is the equivalent length of the combustion chamber.

[0068] Determine whether the following relationship is satisfied: in, is a constant, typically selected based on the specific configuration, size, and power rating of the engine compressor. This equation states that the P2 condition holds true when the primary periodic frequency of the compressor outlet pressure dynamics is close to the Helmholtz frequency.

[0069] P3 conditions: Condition P3 is considered met if the engine is heard to emit a low-frequency humming sound that is clearly different from normal operation, or a continuous burst of explosions, or if flames are observed at the engine outlet.

[0070] During the test monitoring process, if the collected parameters meet any one of P1, P2, and P3, it is determined that the engine has reached the surge limit, and the process proceeds to step 750. If any two of P1, P2, and P3 are not met, the process returns to step 720 and continues to reduce the variable geometry turbine angle until the engine compressor successfully enters the surge state.

[0071] Step 760: Collect surge boundary data and perform surge boundary measurement tests at other speeds.

[0072] After determining that the engine has successfully entered surge, quickly reduce the fuel flow to its initial setting while increasing the variable turbine geometry angle until it reaches the original design angle (0°). If the engine still meets any of the aforementioned P1, P2, and P3 conditions, continue increasing the variable turbine geometry angle while reducing the fuel injection rate to the idle fuel flow until the engine no longer meets the P1, P2, and P3 conditions. Be careful not to shut off the fuel pump directly, as this can damage the engine.

[0073] In this method, a variable-geometry turbine replaces part of the fuel step injection function. Together, they shift the compressor operating point to the aerodynamic instability boundary. The fuel step flow rate replaced by the variable-geometry turbine reduces the combustion chamber outlet gas temperature, thereby reducing the turbine thermal load caused by the fuel step while maintaining the basic aerodynamic characteristics of the fuel step impact on the compressor. The portion of the fuel step replaced by the variable-geometry turbine can also be converted using some of the formulas in step 720 to facilitate standardized comparative analysis.

[0074] The embodiment of the present application also provides a method for determining the aerodynamic instability boundary in a whole machine environment. The method adopts a composite method of a variable geometry turbine and high-pressure gas charging, uses high-pressure charging to simulate the dynamic pressure change caused by the fuel step, and thus completely replaces the fuel step. The gasping method combining high-pressure gas charging with a variable geometry turbine can further reduce the thermal load of the turbine during the gasping process. At the same time, compared with the fuel step gasping method, the high-pressure gas charging method is more flexible and accurate. Figure 12 As shown, the specific process includes: Step 800: Place the engine on the stage and set it to the initial position.

[0075] Install the modified engine on the test bench and secure the external air reservoir in place. Calibrate the initial position of the variable geometry turbine. Based on the original engine geometry, calibrate the modified variable geometry turbine stator to 0° so that the variable geometry turbine angle matches the original geometry. Repeatedly adjust the position of the actuator ring to calibrate the relationship between the actuator position and the variable geometry turbine blade angle. Set the calibrated variable geometry turbine angle to 0°. Open the air reservoir valve to allow air to be compressed by the compressor and enter the air reservoir.

[0076] Step 810: Set the engine to a test speed.

[0077] Ensure the variable turbine geometry is set to 0° and start the engine. Use the pre-modified engine starting control scheme to ignite the engine and bring it to idle speed. Maintaining the variable turbine geometry, gradually accelerate the engine using the pre-modified engine acceleration control scheme until the engine reaches the desired speed for this test.

[0078] Step 820: Adjust the variable geometry turbine angle to reduce the engine throat area.

[0079] Maintain the engine at the speed required for the surge test and gradually reduce the variable geometry turbine angle in 1° increments. While reducing the variable geometry angle, the engine's operating conditions should remain largely unchanged. After each complete adjustment of the variable geometry angle, wait until typical engine performance parameters, such as speed and total pressure ratio, have stabilized before making further adjustments. Importantly, do not force the engine into surge directly through variable geometry adjustments. Surge margin should be maintained to allow for the dynamic process of a sudden pressure increase to be simulated when subsequent surge methods are used.

[0080] Step 830: Dynamically collect compressor outlet pressure parameters as surge judgment criteria.

[0081] A dynamic pressure sensor should be placed at the engine compressor outlet to determine whether the compressor has entered a surge state. Considering that the engine is in normal operating condition during the test, high-temperature combustion gas in the combustion chamber may flow back to the engine end during the surge process, so special consideration should be given to the selection of the dynamic pressure sensor. The surge frequency is approximately 100Hz, and the probe and acquisition equipment frequencies are 20kHz and 200kHz, respectively, which meet the requirements for measuring surge phenomena. The dynamic pressure probe measurement and acquisition system is started synchronously with engine startup. The results of the dynamic pressure acquisition system are monitored at all times.

[0082] Step 840: Close the valve of the air storage chamber and heat the air storage chamber.

[0083] After the engine reaches its designed speed, the pressure inside the reservoir chamber should equal that in the combustion chamber because the reservoir valve is open. At this point, the reservoir valve is closed, isolating the chamber. Subsequently, the reservoir chamber is heated using the heating device within the reservoir chamber. To ensure that the gas within the reservoir chamber can quickly raise the pressure within the combustion chamber during inflation to achieve the desired gasping effect, the reservoir chamber heating temperature should at least meet the following requirements: Where, is the temperature of the gas storage chamber after heating, is the combustion chamber inlet gas temperature, is the combustion chamber inlet pressure, The inlet pressure of the combustion chamber when the aerodynamic instability boundary is reached during the gasping process. is the equivalent aerodynamic volume of the combustion chamber, is the aerodynamic volume of the air storage chamber.

[0084] In order to completely replace the fuel with high pressure gas, the step coefficient isc If the fuel step is forced to breathe in, the heating temperature of the air storage chamber should meet the following requirements: Where, is the temperature of the gas storage chamber after heating, is the combustion chamber inlet gas temperature, is the combustion chamber inlet pressure, The inlet pressure of the combustion chamber when the aerodynamic instability boundary is reached during the gasping process. is the equivalent aerodynamic volume of the combustion chamber, is the aerodynamic volume of the air storage chamber, is the calorific value of fuel, is the specific heat capacity of the gas, is the pre-set fuel step coefficient, Less than a , a is a fuel step coefficient capable of causing the compression system to enter a surge state by using a fuel step method alone; is the air flow rate at the compressor outlet, It is the fuel flow rate when the engine is in stable working condition.

[0085] In actual operation, a corresponding dynamic pressure monitoring probe is often added between the inflation / valve and the high-pressure air chamber. During the heating of the high-pressure air chamber and the inflation process, the pressure measured by the probe should be kept at Above.

[0086] The heating time of the air reservoir depends on the volume of the air reservoir and the performance of the heating equipment. If the heating time is too long, you can stop the engine first, wait for the air reservoir to heat up, then start the engine and accelerate it back to the test speed before continuing the test.

[0087] At low test speeds, the engine can be accelerated to the design speed to build up a high pressure in the air reservoir, and then the speed can be reduced to a lower test speed for deflation. This can reduce the heating time or even eliminate the heating process.

[0088] Step 850: Deflate the air storage chamber to simulate a fuel step. After heating the reservoir, maintain the variable turbine geometry and adjust the valve to bleed air. The specific bleed procedure and bleed hole selection should be determined based on the specific test plan. Bleeding the reservoir will cause the compressor to enter a surge state.

[0089] Step 860: Determine if the engine is breathing heavily according to the following conditions: If at least one of P1, P2, and P3 is met, the engine is determined to have entered a surge state, and the process proceeds to step 870. If none of P1, P2, and P3 is met, the process returns to step 820 to continue reducing the variable geometry turbine angle, or returns to step 840 to further heat the air reservoir before continuing to charge and force the engine to surge. This cycle continues until the engine enters a surge state.

[0090] When the engine enters the surge state, the variable geometry turbine angle should be quickly opened, the engine fuel supply should be reduced and the charging valve should be closed after the data collection is completed, so that the engine can stop surge.

[0091] Step 870: Collect surge boundary data and perform surge boundary measurement tests at other speeds.

[0092] In this approach, a variable-geometry turbine and high-pressure gas charging partially replace the fuel step injection function. Together, they shift the compressor operating point to the aerodynamic instability boundary. This further reduces the thermal load on turbine components during the surge, while maintaining the fundamental aerodynamic process.

[0093] The following is a brief description of the basic process of using the above-mentioned combined method of using a variable geometry turbine and charging high-pressure gas, taking a small turbojet engine equipped with a first-stage centrifugal compressor as an example.

[0094] S1. First, assemble the engine and install the variable-geometry turbine. Modify the engine casing to connect to the air reservoir. Adjust the variable-geometry turbine to zero angle. Open the air reservoir valve.

[0095] S2. Ignite the engine and accelerate according to the normal test process until the engine speed reaches the required speed for the breath test. Maintain the engine speed unchanged.

[0096] S3. Gradually reduce the variable geometry turbine angle in given steps. During this process, the engine control system automatically adjusts the fuel injection rate to ensure that the engine speed does not change.

[0097] S4, close the air storage chamber valve, heat the air storage chamber to a set value, then open the air storage chamber valve to force the engine to breathe.

[0098] S5. Collect dynamic pressure sensor parameters at the inlet and outlet of the engine compressor and filter the collected pressure sensor parameters.

[0099] S6. The engine operating status is determined based on criteria P1, P2, and P3. If the engine is experiencing engine gasps, the compressor parameters obtained before the variable geometry turbine is turned down are used as the compression component instability boundary. If the engine is not experiencing engine gasps, the turndown operation is repeated from step S3.

[0100] S7. After completing the surge boundary measurement for one speed line, continue adjusting the engine operating state, exit the surge state, and accelerate to the second speed line to be measured. Repeat steps S3-S6 until the surge boundaries of all planned speed lines have been measured.

[0101] At this point, the experimental measurement of the aerodynamic instability boundary of the entire engine under this operating condition has been completed. The instability boundary under other operating conditions will be calculated using the above method.

[0102] In the method for determining the aerodynamic instability boundary in the whole machine environment provided in the embodiment of the present application, the two combined surge methods proposed can both take into account the transient effect of the pressure surge, and can better approximate the actual situation of surge caused by rapid acceleration in the real engine working environment. Compared with the fuel step surge that can also take into account the transient effect of the pressure surge, the fuel step-variable geometry turbine compound surge method provided in the embodiment of the present application can use a variable geometry turbine to replace part of the fuel step function, thereby reducing the injected fuel flow, thereby reducing the turbine thermal load during the surge, and ultimately reducing the test risk and cost. Compared with the traditional high-pressure gas charging surge method that also does not cause additional thermal load. The high-pressure gas charging-variable geometry turbine compound surge method provided in the embodiment of the present application can achieve self-sufficiency of the high-pressure gas source by simply modifying and expanding the test engine without providing a high-temperature gas source and without large-scale modification of the test bench, and reduces the test cost and shortens the test preparation time.

[0103] Corresponding to the above-mentioned method for determining the aerodynamic instability boundary under the whole machine environment, the embodiment of the present application also provides an engine test device, such as Figure 13 The engine shown includes: an engine body 900, a modified turbine 910 for testing with a variable flow cross-sectional area, and a test control device 920, wherein the test control device 920 includes: a memory and a processor, wherein the memory is configured to store an executable program; The processor is configured to read and execute the executable program to perform the method for determining the aerodynamic instability boundary in the whole machine environment described in any of the above embodiments.

[0104] In addition to the above-mentioned methods and devices, an embodiment of the present application may also be a computer program product, which includes computer program instructions. When the computer program instructions are executed by a processor, the processor executes the steps of the method for determining the aerodynamic instability boundary in the whole machine environment according to various embodiments of the present application described in any of the above embodiments of this specification.

[0105] The computer program product may be written in any combination of one or more programming languages ​​to implement the program code for performing the operations of the embodiments of the present application, including object-oriented programming languages ​​such as Java, C++, and conventional procedural programming languages ​​such as "C" or similar programming languages. The program code may be executed entirely on the user's computing device, partially on the user's computing device, as a standalone software package, partially on the user's computing device and partially on a remote computing device, or entirely on a remote computing device or server.

[0106] In addition, an embodiment of the present application further provides a storage medium having a computer program stored thereon. When the computer program is executed by a processor, the method for determining the aerodynamic instability boundary in the whole machine environment described in any of the above embodiments is implemented.

[0107] Those skilled in the art will appreciate that all or some of the steps, systems, and functional modules / units in the methods, systems, and devices disclosed above may be implemented as software, firmware, hardware, or any combination thereof. In hardware implementations, the division between functional modules / units described above does not necessarily correspond to the division between physical components; for example, a physical component may have multiple functions, or a function or step may be performed collaboratively by several physical components. Some or all components may be implemented as software executed by a processor, such as a digital signal processor or microprocessor, or as hardware, or as an integrated circuit, such as an application-specific integrated circuit. Such software may be distributed on computer-readable media, which may include computer storage media (or non-transitory media) and communication media (or transient media). As is well known to those skilled in the art, the term "computer storage media" encompasses volatile and non-volatile, removable and non-removable media implemented in any method or technology for storing information, such as computer-readable instructions, data structures, program modules, or other data. Computer storage media includes, but is not limited to, RAM, ROM, EEPROM, flash memory or other memory technology, CD-ROM, digital versatile disks (DVD) or other optical disk storage, magnetic cassettes, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other medium that can be used to store the desired information and can be accessed by a computer. In addition, as is well known to those skilled in the art, communication media typically embodies computer-readable instructions, data structures, program modules, or other data in a modulated data signal such as a carrier wave or other transport mechanism, and may include any information delivery media.

Claims

1. A method for determining the aerodynamic instability boundary in a complete machine environment, characterized in that: Applied to an engine test device, the engine test device includes: an engine body and a modified test turbine with a variable flow cross-sectional area, the engine body includes a compression system, and the method includes: Controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a fuel step method, or controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a high-pressure gas charging method; An instability boundary point of the engine test device is determined, and an aerodynamic instability boundary of the compression system at a current rotational speed is determined based on the instability boundary point.

2. The method according to claim 1, characterized in that In the case where the compression system is controlled to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a fuel step method, the method of controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a fuel step method includes: Rotating the stator blades of the modified turbine for the test to a calibrated initial position; at the initial position, the stator blades of the modified turbine for the test are in the same position as the stator blades of the unmodified turbine; Ignite and start the engine test device, and accelerate it to the set speed for the whole engine breathing test; The stator blades of the modified turbine for the test are controlled to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test, and the compression system is forced to surge in combination with a fuel step method until the compression system enters a surge state.

3. The method according to claim 2, characterized in that The step of controlling the stator blades of the modified turbine for the test to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test, and combining a fuel step method to force the compression system to surge until the compression system enters a surge state, includes: Continuously adjust the fuel step coefficient b So that the fuel supply system can meet the fuel flow according to b The ratio of times changes step by step, according to the final determined fuel step coefficient b Calculate the flow cross-sectional area required to be reached by the modified turbine for the test, and control the stator blades of the modified turbine for the test to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test to the calculated flow cross-sectional area; wherein, the fuel step coefficient adjusted at any time b Less than the fuel step coefficient a , a is a fuel step coefficient capable of causing the compression system to enter a surge state by using a fuel step method alone; After the flow cross-sectional area of ​​the modified turbine for the test reaches the calculated flow cross-sectional area, the fuel step coefficient finally determined is used. b Performing surge testing and determining whether the compression system enters a surge state; In response to the result of determining that the compression system has not entered a surge state, the stator blades of the modified turbine for the test are controlled to rotate gradually to continue reducing the flow cross-sectional area, and after each reduction in the flow cross-sectional area, the fuel step coefficient finally determined is used. b A fuel step is performed and it is determined whether the compression system enters a surge state until the compression system enters a surge state.

4. The method according to claim 3, characterized in that The fuel step coefficient is continuously adjusted b To make the fuel flow rate follow b The ratio of the times has a step change, including: Set the initial fuel step coefficient for fuel flow b and set the initial fuel step coefficient b As the current fuel step coefficient b , perform the following fuel step coefficient adjustment operation until the fuel flow rate is in accordance with b A step change in the ratio of times: The oil step coefficient adjustment operation includes: According to the current fuel step coefficient b adjusting the flow cross-sectional area of ​​the modified turbine for the test and increasing the amount of fuel supplied to the combustion chamber; Determine whether the oil supply system can meet the requirements under the adjusted flow cross-sectional area. b The fuel flow rate can be changed in steps by a factor of 2. Inability to follow b When the fuel flow rate step change is achieved at a ratio of times, continue to adjust the fuel step coefficient b and adjust the fuel step coefficient b As updated current fuel step factor b , continue the fuel step coefficient adjustment operation.

5. The method according to claim 1, wherein In the case where the compression system is controlled to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining a high-pressure gas charging method, the engine testing device further comprises: an air storage chamber, a heating device, the compression system comprises: a compressor, and the engine body comprises: a combustion chamber; The method of controlling the compression system to enter a surge state by adjusting the flow cross-sectional area of ​​the modified turbine for testing and combining the method of charging high-pressure gas comprises: Rotating the stator blades of the modified turbine for the test to a calibrated initial position; at the initial position, the stator blades of the modified turbine for the test are in the same position as the stator blades of the unmodified turbine; Ignite and start the engine test device, and accelerate it to the set speed for the whole engine breathing test; The stator blades of the modified turbine for the test are controlled to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test, and combined with the high-pressure gas charging method, the compression system is forced to surge until the compression system enters a surge state; wherein, the high-pressure gas charging method is a method of using the heating device to heat the high-pressure gas from the compressor in the air storage chamber and output the high-heat and high-pressure gas to the combustion chamber.

6. The method according to claim 5, characterized in that The outer wall of the flow channel at the compressor outlet is provided with a plurality of vent holes, the air storage cavity is connected to the vent holes by a vent pipe, and the vent pipe is provided with a valve; The method comprises: controlling the stator blades of the modified turbine for the test to rotate gradually to reduce the flow cross-sectional area of ​​the modified turbine for the test, and combining the high-pressure gas charging method to force the compression system to surge until the compression system enters a surge state, comprising: Setting an initial flow cross-sectional area required to be achieved by the modified turbine for the test, setting an initial heating temperature for the heating device to heat the air storage chamber, using the set initial flow cross-sectional area as the current flow cross-sectional area, using the set initial heating temperature as the current heating temperature, and performing the following surge operation until the compression system enters a surge state; The breathing operation includes: closing the valve, and using the heating device to heat the high-pressure gas in the gas storage chamber at the current heating temperature to generate high-heat high-pressure gas; opening a valve to allow the high-temperature and high-pressure gas in the gas storage chamber to be filled into the combustion chamber, thereby forcing the compression system to surge, and determining whether the compression system enters a surge state during the surge process; In response to the judgment result that the compression system has not entered the surge state, the current flow cross-sectional area is reduced to obtain an updated current flow cross-sectional area and / or the current heating temperature is increased to obtain an updated current heating temperature, and the surge operation is continued using the updated current flow cross-sectional area and / or the updated current heating temperature.

7. The method according to claim 6, characterized in that The minimum area of ​​a single vent hole is as follows: ； in, is the minimum area of ​​a single vent, is the outlet mass flow rate when the compressor is working stably at the test working point, is the static temperature at the outlet of the compressor when it is working stably at the test working point, is the static pressure at the outlet of the compressor when it is working stably at the test working point, is the correction coefficient, which is determined according to the selection of valve and vent. is the number of vents, is the specific heat ratio, is the gas constant, is the temperature of the gas in the gas storage chamber.

8. The method according to claim 5, characterized in that The heating temperature of the heating device as follows: and ; ； in, is the temperature of the gas storage chamber after heating, is the combustion chamber inlet gas temperature, is the combustion chamber inlet pressure, is the inlet pressure of the combustion chamber when it reaches the aerodynamic instability boundary during the gasping process. is the equivalent aerodynamic volume of the combustion chamber, is the aerodynamic volume of the air storage chamber, is the calorific value of fuel, is the specific heat capacity of the gas, is the pre-set fuel step coefficient, Less than a , a is a fuel step coefficient capable of causing the compression system to enter a surge state by using a fuel step method alone; is the air flow rate at the compressor outlet, It is the fuel flow rate when the engine is in stable working condition.

9. The method according to claim 3, characterized in that Whether the engine enters the surge state is determined based on whether the operating information of the engine test device satisfies a determination condition for entering the surge state; The operation information of the engine test device includes: a physical quantity used to characterize the performance of the compression system, and a signal used to characterize the operating state of the engine test device; The judgment condition for entering the surge state includes at least one of the following: a physical quantity used to characterize the performance of the compression system meets a set surge condition, and a signal used to characterize the operating state of the engine test device shows a set surge signal.

10. An engine test device, characterized in that: include: Engine body, modified turbine with variable flow cross-sectional area for testing and test control device, The test control device includes: a memory and a processor, wherein the memory is configured to store an executable program; The processor is configured to read and execute the executable program to perform the method for determining the aerodynamic instability boundary in a whole machine environment according to any one of claims 1 to 9.

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