An experimental device and method for the surge boundary of an aero-engine compressor

Abstract

This application belongs to the technical field of determining the surge boundary of an aero-engine compressor, and particularly relates to an aero-engine compressor surge boundary test device and method. Among them, the aero-engine compressor surge boundary test device includes: a fuel tank; a high-pressure fuel tank; a piston cylinder; two pistons connected by a connecting rod and arranged inside the piston cylinder, with an oil suction control cavity formed therebetween. Among them, an oil passage cavity is formed between one piston and the corresponding end of the piston cylinder, and an oil injection control cavity is formed between the other piston and the corresponding end of the piston cylinder; the oil suction control cavity and the oil injection control cavity are connected to the high-pressure fuel tank through pipelines, and a servo valve is arranged on this pipeline; the oil passage cavity is connected to the fuel tank through a pipeline, a through-oil valve is arranged on this pipeline, and it is connected to the fuel injector of the aero-engine combustion chamber through a pipeline; a controller, connected to the servo valve and the through-oil valve.

Description

Technical Field

[0001] This application belongs to the technical field of determining the surge boundary of an aero-engine compressor, and particularly relates to an aero-engine compressor surge boundary test device and method. Background Art

[0002] Currently, the surge boundary of an aero-engine compressor is mostly obtained through compressor component tests. On the component test bench, the component tester keeps the compressor corrected speed unchanged, gradually reduces the exhaust area of the test piece, raises the operating point of the compressor to surge, and simultaneously measures the compressor inlet corrected flow rate and pressure ratio, thereby obtaining a point on the surge boundary. Repeating this process at multiple corrected speeds can obtain the surge boundary of the compressor.

[0003] However, after the compressor is installed on the aero-engine, its surge boundary changes compared to the surge boundary obtained from component tests. To evaluate the stability of the engine, it is necessary to conduct aero-engine stability tests to indirectly evaluate the stability of the aero-engine, but the surge boundary of the compressor cannot be accurately obtained.

[0004] Some researchers also use the method of modifying the fuel supply law of the aero-engine control system to increase the instantaneous fuel supply in the aero-engine combustion chamber to cause the compressor to surge and obtain the compressor surge boundary. However, this technical solution has the following defects:

[0005] Under high corrected speed conditions, to cause the compressor to surge by instantaneously supplying fuel to the aero-engine combustion chamber, the fuel flow rate needs to increase by 1 - 2 times based on the steady-state fuel flow rate. The fuel supply capacity of the aero-engine fuel supply system is limited and it is difficult to meet the requirement of this fuel supply amount, which is not applicable to all states of the compressor;

[0006] The fuel supply speed of the aero-engine fuel supply system is limited and it is difficult to ensure instantaneously supplying fuel to the aero-engine combustion chamber, which will cause the compressor to accelerate. Although the corresponding surge points can be obtained, the adjustable blades and speed of the compressor will change, showing a large deviation from the actual situation.

[0007] In view of the existence of the above technical defects, this application is proposed.

[0008] It should be noted that the disclosure of the above background art content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. Without clear evidence indicating that the above content was publicly available on the filing date of this application, the above background art should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0009] The object of the present application is to provide a surge margin test device and method for an aero-engine compressor to overcome or mitigate at least one aspect of the known technical defects.

[0010] The technical solution of the present application is as follows:

[0011] On the one hand, a surge margin test device for an aero-engine compressor is provided, including:

[0012] A fuel tank;

[0013] A high-pressure fuel tank;

[0014] A piston cylinder, the inner wall of which has an annular protrusion;

[0015] Two pistons, connected by a connecting rod, are arranged in the piston cylinder and connected by the connecting rod passing through the annular protrusion; an oil suction control cavity is formed therebetween. Among them, an oil passage cavity is formed between one piston and the corresponding end of the piston cylinder; an oil injection control cavity is formed between the other piston and the corresponding end of the piston cylinder and an oil suction control cavity is formed between it and the annular protrusion; the oil suction control cavity and the oil injection control cavity are connected to the high-pressure fuel tank through pipelines, and a servo valve is arranged on this pipeline; the oil passage cavity is connected to the fuel tank through a pipeline, a through-flow valve is arranged on this pipeline, and is connected to the fuel injector of the aero-engine combustion chamber through a pipeline.

[0016] A controller, connected to the servo valve and the through-flow valve, when receiving a fuel step command, controls the through-flow valve to open, and controls the servo valve to introduce hydraulic oil from the high-pressure fuel tank into the oil suction control cavity, so that the two pistons move towards the oil injection control cavity direction, so that the required fuel quantity required by the fuel step command is sucked from the fuel tank in the oil passage cavity. Furthermore, the through-flow valve is closed, and the servo valve is controlled to introduce hydraulic oil from the high-pressure fuel tank into the oil injection control cavity, so that the two pistons move towards the oil passage cavity direction, so that the fuel sucked in the oil passage cavity is injected into the combustion chamber through the fuel injector of the combustion chamber according to the time required by the fuel step command.

[0017] Optionally, in the above-mentioned surge margin test device for an aero-engine compressor, there are multiple groups of piston cylinders and their corresponding pistons, and the sizes between each group are different.

[0018] Optionally, in the above-mentioned surge margin test device for an aero-engine compressor, it further includes:

[0019] Sensors related to the acquisition of the compressor inlet corrected flow rate, installed on the aero-engine;

[0020] Sensors related to the acquisition of the compressor pressure ratio, installed on the aero-engine;

[0021] The sensors related to the acquisition of the compressor inlet corrected flow rate and the sensors related to the acquisition of the compressor pressure ratio are connected to the controller, and the compressor inlet corrected flow rate and its pressure ratio are acquired according to the set frequency.

[0022] Optionally, in the above aero-engine compressor surge margin test device, it further includes:

[0023] Fuel quantity acquisition-related sensors are arranged in the piston barrel or the pipeline connecting the oil passage chamber and the combustion chamber injector nozzle, connected to the controller, and the fuel injection quantity from the oil passage chamber into the combustion chamber is acquired at a set frequency.

[0024] Optionally, in the above aero-engine compressor surge margin test device, it further includes:

[0025] A hydraulic oil tank is connected to the high-pressure oil tank through a pipeline, and a hydraulic pump is arranged on this pipeline;

[0026] The controller is connected to the hydraulic pump to be able to control the hydraulic pump to pressurize the hydraulic oil in the hydraulic oil tank and transport it into the high-pressure oil tank, so that the hydraulic oil in the high-pressure oil tank reaches the set value.

[0027] Optionally, in the above aero-engine compressor surge margin test device, the controller is connected to the aero-engine control system to be able to transmit a fuel step command to the aero-engine control system;

[0028] After receiving the step command, the aero-engine control system locks the fuel supply quantity of the aero-engine fuel supply system and the angle of the adjustable compressor blades.

[0029] Optionally, in the above aero-engine compressor surge margin test device, when the aero-engine control system determines that the aero-engine has a surge or over-temperature, it feeds back relevant signals to the controller, and the controller controls the servo valve to stop injecting the hydraulic oil from the high-pressure oil tank into the oil injection control chamber.

[0030] On the other hand, a method for testing the aero-engine compressor surge margin is provided, including:

[0031] Construct an aero-engine transient model;

[0032] Using the aero-engine transient model, calculate the fuel step form at each corrected speed. The fuel step form includes the step fuel quantity and the step time. When injecting fuel into the combustion chamber according to the fuel step form, when the step fuel quantity reaches the peak value, the change in the compressor speed is less than 0.5%, and the compressor operating point reaches the surge margin;

[0033] Corresponding to each corrected speed, generate a fuel step command according to the fuel step form, and then inject fuel into the aero-engine combustion chamber according to the fuel step command to conduct the aero-engine compressor surge margin test.

[0034] Optionally, in the above method for testing the surge margin of an aero-engine compressor, for each corrected speed, among the fuel step forms that meet the requirements, select the fuel step form with a small step fuel quantity and a long step time to conduct the aero-engine compressor surge margin test.

[0035] Optionally, in the above method for testing the surge margin of an aero-engine compressor, it further includes:

[0036] Change the output frequency of the process parameters of the aero-engine transient model, find the minimum output frequency of the process parameters that can output enough points within the step time to describe the upward movement process of the compressor operating point, and collect the corrected inlet flow rate and pressure ratio data of the compressor at this output frequency of the process parameters when conducting the aero-engine compressor surge margin test. Description of the Drawings

[0037] Figure 1 It is a schematic diagram showing that increasing the fuel flow rate will accelerate the aero-engine and raise the operating point of the compressor provided by the embodiment of the present application;

[0038] Figure 2 It is a schematic diagram of the fuel step form provided by the embodiment of the present application;

[0039] Figure 3 It is a schematic diagram of the fuel step forms with different step times provided by the embodiment of the present application;

[0040] Figure 4 It is a schematic diagram of the calculation results of the fuel step forms with different step times provided by the embodiment of the present application;

[0041] Figure 5 It is a schematic diagram of the fuel step forms with different step fuel quantities provided by the embodiment of the present application;

[0042] Figure 6 It is a schematic diagram of the calculation results of the fuel step forms with different step fuel quantities provided by the embodiment of the present application;

[0043] Figure 7 It is a schematic diagram of the fuel step times corresponding to different compressor rotor moments of inertia provided by the embodiment of the present application;

[0044] Figure 8 It is a schematic diagram of the operating point trajectory of the compressor at output frequencies of 20 Hz, 50 Hz, 100 Hz, and 200 Hz provided by the embodiment of the present application;

[0045] Figure 9 It is a schematic diagram of the aero-engine compressor surge margin test device provided by the embodiment of the present application;

[0046] Wherein:

[0047] 1 - Fuel tank; 2 - High - pressure fuel tank; 3 - Piston cylinder; 4 - Two pistons; 5 - Servo valve; 6 - Oil - passing valve; 7 - Aero - engine; 8 - Controller; 9 - Sensors related to collecting compressor inlet conversion flow; 10 - Sensors related to collecting compressor pressure ratio; 11 - Sensors related to collecting fuel quantity; 12 - Hydraulic oil tank; 13 - Hydraulic pump; 14 - Aero - engine control system; 15 - Aero - engine fuel supply system.

[0048] To better illustrate this embodiment, some components in the drawings are omitted, enlarged or reduced, which do not represent the dimensions of the actual product. In addition, the drawings are only for illustrative purposes and cannot be construed as a limitation of this patent. Detailed implementation manners

[0049] To make the technical solutions and their advantages of this application clearer, the technical solutions of this application will be further described clearly and completely with reference to the drawings. It can be understood that the specific embodiments described herein are only partial embodiments of this application, which are only used to explain this application rather than limit this application. It should be noted that for the convenience of description, only the parts related to this application are shown in the drawings, and other related parts can refer to the general design. Without conflict, the embodiments and the technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0050] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application should be the ordinary meanings understood by those of ordinary skill in the field to which this application belongs. The words indicating directions such as "up", "down", "left", "right", "center", "vertical", "horizontal", "inner", "outer", etc. used in the description of this application are only used to represent relative directions or position relationships, rather than implying that the device or component must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative position relationship may also change accordingly, so it cannot be construed as a limitation of this application. The terms "first", "second", "third" and similar terms used in the description of this application are only for descriptive purposes to distinguish different components and cannot be understood as indicating or implying relative importance. The terms "a", "an" or "the" and similar words used in the description of this application should not be understood as an absolute limitation of quantity but should be understood as having at least one. The terms "including" or "comprising" and similar words used in the description of this application are intended to cover the elements or objects appearing before this word and the elements or objects listed after this word and their equivalents, without excluding other elements or objects.

[0051] In addition, it should be noted that, unless otherwise clearly specified and limited, the similar terms such as "installed", "connected", and "linked" used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium, and it can also be the communication inside two components. Those skilled in the art can understand their specific meanings in this application according to the specific circumstances.

[0052] The following will further elaborate on this application in conjunction with the Figures 1 to 9 accompanying drawings.

[0053] When the aero-engine operates at different steady-state speeds n2, increasing the fuel flow rate will accelerate the aero-engine, thereby raising the operating point of the compressor. During the engine acceleration, the fuel flow rate in the combustion chamber gradually increases, the operating line of the high-pressure compressor will rise, and the engine speed increases. If the fuel flow rate is increased rapidly, the operating point of the compressor can be raised, but due to inertia, the rotor of the aero-engine is not yet able to accelerate in time. This will cause the operating point of the compressor to approximately rise along the constant-speed line. When the instantaneous fuel flow rate is high enough, the operating point of the compressor will be raised to the surge boundary, and thus the surge boundary of the engine is obtained, as Figure 1 shown.

[0054] In order to approximately move the operating point of the compressor up to the surge boundary at a constant speed, the step fuel quantity and the step time need to be coordinated with each other. One of the important problems to be solved in this application is how to determine the step fuel quantity and the step time.

[0055] Construct an aero-engine transient model. Assume that when n2 = 85% for a certain aero-engine, determine the fuel step scheme at this speed, calculate the fuel step forms for different step times, and the obtained results are as Figure 4 shown. Calculate the fuel step forms for different step fuel quantities, and the obtained results are as Figure 6 shown. Through this calculation, the fuel step form at the speed of n2 = 85% can be determined.

[0056] There are two criteria for the fuel step forms at each speed of the aero-engine:

[0057] When the fuel step reaches the peak, the change in speed is less than 0.5%;

[0058] When the fuel step reaches the peak, the operating point of the compressor reaches the surge boundary.

[0059] If both of these are satisfied, the fuel step form can be determined.

[0060] During a fuel step, the sensitivity of the compressor rotor to fuel changes can determine the fuel step time, which is mainly related to the moment of inertia of the compressor rotor. The fuel step times corresponding to different compressor rotor moments of inertia are as Figure 7 shown. For example, when the moment of inertia of the compressor rotor is 10 kg·m2, the fuel step time can be directly determined to be 0.25 s through Figure 5 .

[0061] For each corrected speed, in the fuel step form that meets the requirements, a fuel step form with a small step fuel quantity and a long step time should be selected as much as possible for the aeroengine compressor surge margin test to reduce the hardware requirements for the equipment.

[0062] The purpose of the fuel step is to obtain the surge margin of the compressor in the whole-machine environment. The quantities describing the compressor surge margin are the corrected flow rate at the compressor inlet and the pressure ratio. This requires these two quantities to be measured at the moment of the fuel step. However, the main problem faced is the determination of the acquisition frequency.

[0063] As mentioned above, the time of the fuel step is about 0.1 s, which requires high-frequency acquisition. However, if the required frequency is too high, it will increase the cost of the test equipment. Therefore, this application will give a simulation-based acquisition frequency evaluation method.

[0064] The calculation of the aeroengine transient model requires a time step for calculating the transient state, which is used to calculate the engine parameters in the next cycle. In the aeroengine transient model, this time step is generally set to 0.001 s. During the study of the fuel step test process, the minimum parameter acquisition frequency required by the aeroengine during the fuel step can be evaluated through transient simulation. The method is as follows:

[0065] Set the output frequency of the transient process parameters. Generally, when studying the working state of the engine process, an output frequency of 10 Hz meets the requirements. The calculation step is 0.001 s, that is, every 100 steps are calculated, and 1 point is output every 0.1 s;

[0066] Check the trajectory of the compressor operating point in the compressor characteristic diagram during the fuel step process. If the trajectory points meet the data analysis requirements, it is considered that the output frequency is appropriate and can be used as the data acquisition frequency for the aeroengine compressor surge margin test.

[0067] The trajectory conditions of the compressor operating point at output frequencies of 20 Hz, 50 Hz, 100 Hz, and 200 Hz are as Figure 8As shown, at output frequencies of 20 Hz and 50 Hz, only 2 to 4 points can be collected at the moment of fuel step change, which cannot meet the requirements for data analysis during the fuel step process. In the case of 100 Hz to 200 Hz, 10 to 20 points can be collected at the moment of fuel step change, basically describing the trajectory of the upward movement of the compressor operating point. Therefore, when conducting the compressor surge boundary test of an aeroengine and during fuel step change, the data acquisition frequency can be set to 100 Hz to 200 Hz. Under the condition of meeting the analysis accuracy requirements, 100 Hz can be selected.

[0068] In order to reliably implement the compressor surge boundary test of an aeroengine, the present application provides an aeroengine compressor surge boundary test device, as Figure 9 shown, including:

[0069] Fuel tank 1;

[0070] High-pressure fuel tank 2;

[0071] Piston cylinder 3, the inner wall of which has an annular protrusion;

[0072] Two pistons 4, arranged in the piston cylinder 3 and connected by a connecting rod passing through the annular protrusion; wherein, an oil passage chamber is formed between one piston 4 and the corresponding end of the piston cylinder 3; an oil injection control chamber is formed between the other piston 4 and the corresponding end of the piston cylinder 3, and an oil suction control chamber is formed between the other piston 4 and the annular protrusion; the oil suction control chamber and the oil injection control chamber are connected to the high-pressure fuel tank 2 through pipelines, and a servo valve 5 is arranged on this pipeline; the oil passage chamber is connected to the fuel tank 1 through a pipeline, a through-oil valve 6 is arranged on this pipeline, and is connected to the fuel injector of the combustion chamber of the aeroengine 7 through a pipeline.

[0073] A controller 8, connected to the servo valve 5 and the through-oil valve 6. When receiving a fuel step instruction, it controls the through-oil valve 6 to open, and controls the servo valve 5 to introduce hydraulic oil from the high-pressure fuel tank 2 into the oil suction control chamber, so that the two pistons 4 move towards the oil injection control chamber direction, so that the fuel tank 1 sucks the fuel quantity required by the fuel step instruction in the oil passage chamber. Furthermore, it closes the through-oil valve 6, controls the servo valve 5 to introduce hydraulic oil from the high-pressure fuel tank 2 into the oil injection control chamber, so that the two pistons 4 move towards the oil passage chamber direction, so that the fuel sucked in the oil passage chamber is injected into the combustion chamber through the fuel injector of the combustion chamber according to the time required by the fuel step instruction, to make up for the limitations of the limited fuel supply capacity and limited fuel supply speed of the fuel supply system 7 of the aeroengine, and ensure that reliable compressor surge boundary results can be obtained within the full state range of the compressor.

[0074] Further, in the above-described aero-engine compressor surge margin test device, there are multiple piston cylinders 3 and their corresponding pistons 4, and the sizes of each group are different. Furthermore, different specifications of piston cylinders 3 and their corresponding pistons 4 can be selected for operation according to the fuel quantity and time required by the actual fuel step command, which can be conveniently and flexibly adapted to different situations.

[0075] Further, in the above-described aero-engine compressor surge margin test device, there are two annular protrusions, and an inflation chamber is formed therebetween. Gas is injected into the inflation chamber to prevent hydraulic oil from entering the fuel.

[0076] Further, in the above-described aero-engine compressor surge margin test device, it further includes:

[0077] Relevant sensors 9 for collecting the compressor inlet corrected flow rate, which are installed on the aero-engine 7;

[0078] Relevant sensors 10 for collecting the compressor pressure ratio, which are installed on the aero-engine 7;

[0079] The relevant sensors 9 for collecting the compressor inlet corrected flow rate and the relevant sensors 10 for collecting the compressor pressure ratio are connected to the controller 8, and the compressor inlet corrected flow rate and its pressure ratio are collected at a set frequency.

[0080] Further, in the above-described aero-engine compressor surge margin test device, it further includes:

[0081] Relevant sensors 11 for collecting the fuel quantity, which are arranged on the piston cylinder 3 or the pipeline connecting the oil passage chamber and the combustion chamber fuel injector, and are connected to the controller 8, and the fuel injection quantity from the oil passage chamber into the combustion chamber is collected at a set frequency.

[0082] Further, in the above-described aero-engine compressor surge margin test device, it further includes:

[0083] A hydraulic oil tank 12 is connected to the high-pressure oil tank 2 through a pipeline, and a hydraulic pump 13 is arranged on this pipeline;

[0084] The controller 8 is connected to the hydraulic pump 13 to be able to control the hydraulic pump 13 to pressurize the hydraulic oil in the hydraulic oil tank 12 and transport it into the high-pressure oil tank 2 so that the hydraulic oil in the high-pressure oil tank 2 reaches the set value.

[0085] Further, in the above-described aero-engine compressor surge margin test device, the controller 8 is connected to the aero-engine control system 14 to be able to transmit a fuel step command to the aero-engine control system 14;

[0086] After receiving the step command, the aero-engine control system 14 locks the fuel supply quantity of the aero-engine fuel supply system 15 and the angle of the compressor adjustable vane to ensure the accuracy of the test results.

[0087] Further, in the above-described aero-engine compressor surge margin test device, when the aero-engine control system 14 determines that the aero-engine 7 surges or overheats, it feeds back relevant signals to the controller 8, and the controller 8 controls the servo valve 5 to stop introducing the hydraulic oil from the high-pressure oil tank 2 into the fuel injection control chamber, that is, to terminate the fuel step process to ensure safety.

[0088] The various embodiments in the specification are described in a progressive manner. Each embodiment focuses on the differences from other embodiments. For the same or similar parts among the various embodiments, reference can be made to each other.

[0089] So far, the technical solutions of the present application have been described in conjunction with the preferred embodiments shown in the drawings. Those skilled in the art should understand that the protection scope of the present application is obviously not limited to these specific embodiments. Without departing from the principle of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the protection scope of the present application.

Claims

1. An aero-engine compressor surge margin test device, characterized in that, Comprising: Fuel tank (1); High-pressure fuel tank (2); Piston cylinder (3), the inner wall of which has an annular protrusion; Two pistons (4), arranged within the piston cylinder (3) and connected by a connecting rod passing through the annular protrusion; wherein, an oil passage chamber is formed between one piston (4) and the corresponding end of the piston cylinder (3); an oil injection control chamber is formed between the other piston (4) and the corresponding end of the piston cylinder (3), and an oil suction control chamber is formed between the other piston (4) and the annular protrusion; the oil suction control chamber and the oil injection control chamber are connected to the high-pressure fuel tank (2) through pipelines, and a servo valve (5) is provided on this pipeline; the oil passage chamber is connected to the fuel tank (1) through a pipeline, a through-oil valve (6) is provided on this pipeline, and is connected to the fuel injector of the combustion chamber of the aero-engine (7) through a pipeline; A controller (8), connected to the servo valve (5) and the through-oil valve (6), when receiving a fuel step command, controls the through-oil valve (6) to open, and controls the servo valve (5) to introduce hydraulic oil from the high-pressure fuel tank (2) into the oil suction control chamber, causing the two pistons (4) to move towards the oil injection control chamber direction, so that the fuel amount required by the fuel step command is sucked from the fuel tank (1) into the oil passage chamber. Furthermore, the through-oil valve (6) is closed, and the servo valve (5) is controlled to introduce hydraulic oil from the high-pressure fuel tank (2) into the oil injection control chamber, causing the two pistons (4) to move towards the oil passage chamber direction, so that the fuel sucked into the oil passage chamber is injected into the combustion chamber through the fuel injector of the combustion chamber according to the time required by the fuel step command.

2. The aero-engine compressor surge margin test device according to claim 1, characterized in that, There are multiple groups of the piston cylinder (3) and its corresponding piston (4), and the sizes of each group are different.

3. The aero-engine compressor surge margin test device according to claim 1, characterized in that, Further comprising: Relevant sensors (9) for collecting the compressor inlet conversion flow rate, installed on the aero-engine (7); Relevant sensors (10) for collecting the compressor pressure ratio, installed on the aero-engine (7); The relevant sensors (9) for collecting the compressor inlet conversion flow rate and the relevant sensors (10) for collecting the compressor pressure ratio are connected to the controller (8), and the compressor inlet conversion flow rate and its pressure ratio are collected at a set frequency.

4. The aero-engine compressor surge margin test device according to claim 1, characterized in that, Further comprising: Relevant sensors (11) for collecting the fuel amount, arranged on the piston cylinder (3) or the pipeline connecting the oil passage chamber and the fuel injector of the combustion chamber, connected to the controller (8), and the fuel injection amount from the oil passage chamber into the combustion chamber is collected at a set frequency.

5. The aero-engine compressor surge margin test device according to claim 1, characterized in that, Further comprising: Hydraulic oil tank (12), connected to the high-pressure fuel tank (2) through a pipeline, and a hydraulic pump (13) is provided on this pipeline; The controller (8) is connected to the hydraulic pump (13) to be able to control the hydraulic pump (13) to pressurize the hydraulic oil in the hydraulic oil tank (12) and transport it into the high-pressure fuel tank (2) so that the hydraulic oil in the high-pressure fuel tank (2) reaches the set value.

6. The aero-engine compressor surge margin test device according to claim 1, characterized in that, The controller (8) is connected to the aero-engine control system (14) to be able to transmit a fuel step command to the aero-engine control system (14); After receiving the step command, the aero-engine control system (14) locks the fuel supply amount and the angle of the adjustable compressor blades of the aero-engine fuel supply system (15).

7. The aero-engine compressor surge margin test device according to claim 6, characterized in that, When the aero-engine control system (14) determines that the aero-engine (7) has surge or over-temperature, it feeds back relevant signals to the controller (8), and the controller (8) controls the servo valve (5) to stop supplying the hydraulic oil from the high-pressure oil tank (2) into the fuel injection control chamber.

8. An aero-engine compressor surge margin test method, implemented based on the aero-engine compressor surge margin test device according to claim 1, characterized in that, It includes: Constructing an aero-engine transient model; Using the aero-engine transient model to calculate the fuel step forms at various corrected speeds. The fuel step forms include step fuel quantity and step time. When injecting fuel into the combustion chamber according to the fuel step form, when the step fuel quantity reaches the peak value, the change in the compressor speed is less than 0.5%, and the compressor operating point reaches the surge boundary; Corresponding to each corrected speed, generating a fuel step command according to the fuel step form, and then injecting fuel into the combustion chamber of the aero-engine (7) according to the fuel step command to conduct the aero-engine compressor surge boundary test.

9. The method for testing the surge margin of an aero-engine compressor according to claim 8, wherein For each corrected speed, among the fuel step forms that meet the requirements, select the fuel step form with a small step fuel quantity and a long step time to conduct the aero-engine compressor surge boundary test.

10. The method for testing the surge margin of an aero-engine compressor according to claim 8, wherein It also includes: Changing the output frequency of the process parameters of the aero-engine transient model, finding the minimum process parameter output frequency that can output enough points within the step time to describe the upward movement process of the compressor operating point, and collecting the corrected flow rate at the compressor inlet and its pressure ratio data at this process parameter output frequency when conducting the aero-engine compressor surge boundary test.

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