A method for obtaining compressor stability boundary under the condition of an aero-engine as a whole

Abstract

The present application belongs to the technical field of obtaining the compressor stability boundary under the condition of an aircraft engine as a whole, and specifically relates to a method for obtaining the compressor stability boundary under the condition of an aircraft engine as a whole, comprising: making the aircraft engine operate stably at a given compressor conversion speed; increasing the fuel supply of the aircraft engine in a stepwise manner to cause the aircraft engine to surge, and obtaining the stability boundary under the given compressor conversion speed; changing the given compressor conversion speed, repeating the above steps, and obtaining the corresponding stability boundary.

Description

Technical Field

[0001] The present application belongs to the technical field of compressor stability boundary acquisition under the condition of an aircraft engine as a whole, and specifically relates to a method for acquiring the compressor stability boundary under the condition of an aircraft engine as a whole. Background Art

[0002] The compressor is the core component of an aero-engine. Its operating stability directly affects the operating stability of the entire aero-engine. Obtaining the stability boundary of the compressor under the conditions of the entire aero-engine is of great significance for the stable operation design of the aero-engine.

[0003] Currently, the stability boundary of the compressor is mostly obtained by the principle of moving the working line upward during the acceleration process under the condition of the entire aircraft engine, such as Figure 1 As shown, this technical solution has the following defects:

[0004] The converted speed n2r25 of the compressor is calculated through the physical speed n2 of the compressor and the total temperature T25t at the compressor inlet. During the acceleration process of the aircraft engine, when surge occurs, the aircraft engine itself accelerates, and the total temperature T25t at the compressor inlet changes dynamically. Due to the measurement lag of the total temperature sensor, it is difficult to obtain the accurate value of the total temperature T25t at the compressor inlet in real time. The stability boundary of the compressor at each converted speed n2r25 cannot be accurately obtained, and it is difficult to obtain the accurate working margin of the compressor at each converted speed n2r25. At the same time, it is easily affected by changes in the aerodynamic parameters of the aircraft engine, and repeated debugging is required. The process is cumbersome, inefficient, and has a long cycle.

[0005] This application is proposed in view of the above-mentioned technical defects.

[0006] It should be noted that the disclosure of the above background technology 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. In the absence of clear evidence that the above content has been disclosed on the filing date of this application, the above background technology should not be used to evaluate the novelty and creativity of this application. Summary of the Invention

[0007] The purpose of this application is to provide a method for obtaining the compressor stability boundary under the condition of an aircraft engine as a whole, so as to overcome or alleviate at least one of the technical defects of the known ones.

[0008] The technical solution of this application is:

[0009] A method for obtaining a compressor stability boundary under the condition of an aircraft engine as a whole, comprising:

[0010] Make the aircraft engine run stably at a given compressor conversion speed;

[0011] The fuel supply of the aircraft engine is increased stepwise to cause the aircraft engine to surge, and the stability boundary at a given compressor conversion speed is obtained;

[0012] Change the given compressed air conversion speed and repeat the above steps to obtain the corresponding stability boundary.

[0013] According to at least one embodiment of the present application, in the above-mentioned method for obtaining the compressor stability boundary under the condition of the entire aircraft engine, the fuel supply of the aircraft engine is increased in a stepwise manner to cause the aircraft engine to surge, and the stability boundary at a given compressor converted speed is obtained, specifically:

[0014] Based on the stability margin designed for a given compressor converted speed, the fuel supply of the aircraft engine is increased in a step-wise manner. When the aircraft engine does not experience surge, the value of the step-wise increase in fuel supply is increased until the aircraft engine no longer experiences surge, thereby obtaining the stability margin at the given compressor converted speed.

[0015] According to at least one embodiment of the present application, in the above-mentioned method for obtaining the compressor stability boundary under the condition of the entire aircraft engine, the period of step-by-step increasing the fuel supply amount of the aircraft engine is

[0016] in,

[0017] ΔWFJY is the value of the step-wise increase in fuel supply;

[0018] tHM is the actuation rate of the fuel supply valve mechanism of the aircraft engine control system.

[0019] According to at least one embodiment of the present application, in the above-mentioned method for obtaining the compressor stability boundary under the condition of the entire aircraft engine, after the aircraft engine surges, the step-by-step increase in fuel supply is stopped and the aircraft engine surge relief plan is executed.

[0020] According to at least one embodiment of the present application, in the above-mentioned method for obtaining the compressor stability boundary under the conditions of the entire aircraft engine, the flow rate and pressure ratio between the corresponding points on the working line and the points on the stability boundary at each given compressor conversion speed are obtained by interpolating the characteristics of the compressor components.

[0021] This application has at least the following beneficial technical effects:

[0022] Provided is a method for obtaining a compressor stability boundary under the condition of an entire aircraft engine. The method is designed to obtain the stability boundary at a given compressor conversion speed by step-wise increasing the fuel supply of the aircraft engine stably operating at a given compressor conversion speed, causing the aircraft engine to surge. The accuracy of the corresponding relationship between each compressor conversion speed and the stability boundary can be ensured, and the operating margin of the compressor at each conversion speed can be accurately obtained. The process is not easily disturbed by changes in the aerodynamic parameters of the aircraft engine, is highly efficient, and can shorten the cycle. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 This is a schematic diagram of the method for obtaining the compressor stability boundary under existing aircraft engine complete machine conditions;

[0024] Figure 2 Schematic diagram of a method for obtaining a compressor stability boundary under the condition of an entire aircraft engine provided by an embodiment of the present application;

[0025] Figure 3 This is a flow chart of a method for obtaining the compressor stability boundary under the condition of an aircraft engine as a whole provided in an embodiment of the present application.

[0026] In order to better illustrate this embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product. In addition, the drawings are only used for illustrative purposes and should not be understood as limitations on this patent. DETAILED DESCRIPTION

[0027] To make the technical solution and its advantages of the present application clearer, the technical solution of the present application will be described in further detail below in conjunction with the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of the present application and are only used to explain the present application, not to limit the present application. It should be noted that, for ease of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments of the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0028] In addition, unless otherwise defined, the technical or scientific terms used in the description of this application should have the ordinary meanings understood by those of ordinary skill in the art to which this application belongs. The words "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer" used in the description of this application are only used to indicate relative directions or positional relationships, and do not imply that the device or component must have a specific orientation, be constructed, or operate in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. Therefore, they should not be understood as limitations on this application. The words "first," "second," "third," and similar terms used in the description of this application are used only for descriptive purposes to distinguish different components and should not be understood to indicate or imply relative importance. The words "one," "an," or "the" used in the description of this application should not be understood as absolute limitations on quantity, but should be understood as meaning the presence of at least one. The words "include" or "comprises" used in the description of this application mean that the element or object listed before the word includes the elements or objects listed after the word and their equivalents, but does not exclude other elements or objects.

[0029] In addition, it should be noted that, unless otherwise clearly stipulated and limited, the words "install", "connect", "connect" and similar terms 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, an indirect connection through an intermediate medium, or a connection between two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.

[0030] The following is combined with Figures 1 to 3 This application is described in further detail.

[0031] A method for obtaining a compressor stability boundary under the condition of an aircraft engine as a whole, comprising:

[0032] Make the aircraft engine run stably at a given compressor conversion speed;

[0033] The fuel supply of the aircraft engine is increased stepwise to cause the aircraft engine to surge, and the stability boundary at a given compressor conversion speed is obtained;

[0034] Change the given compressed air conversion speed and repeat the above steps to obtain the corresponding stability boundary.

[0035] As for the method for obtaining the compressor stability boundary under the condition of the entire aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that its design obtains the stability boundary at a given compressor conversion speed by step-wise increasing the fuel supply of the aircraft engine that is stably operating at a given compressor conversion speed, thereby causing the aircraft engine to surge. This ensures the accuracy of the correspondence between each compressor conversion speed and the stability boundary, and further accurately obtains the working margin at each compressor conversion speed. The process is not easily disturbed by changes in the aerodynamic parameters of the aircraft engine, is highly efficient, and can shorten the cycle.

[0036] In some optional embodiments, in the above-mentioned method for obtaining the compressor stability boundary under the condition of the entire aircraft engine, the fuel supply of the aircraft engine is increased in a stepwise manner to cause the aircraft engine to surge, and the stability boundary at a given compressor converted speed is obtained, specifically:

[0037] Based on the stability margin designed for a given compressor converted speed, the aircraft engine's fuel supply is increased in a stepwise manner. That is, the specific amount of the stepwise increase in the aircraft engine's fuel supply is initially set based on the stability margin designed for a given compressor converted speed. When the aircraft engine does not experience surge, the value of the stepwise increase in the fuel supply is increased until the aircraft engine does not experience surge, thereby obtaining the stability boundary for the given compressor converted speed.

[0038] In some optional embodiments, in the above-mentioned method for obtaining the compressor stability boundary under the condition of the entire aircraft engine, the period of stepwise increasing the fuel supply amount of the aircraft engine is

[0039] in,

[0040] ΔWFJY is the value of the step-wise increase in fuel supply;

[0041] tHM is the actuation rate of the fuel supply valve mechanism of the aircraft engine control system.

[0042] In some optional embodiments, in the above-mentioned method for obtaining the compressor stability boundary under the condition of the entire aircraft engine, after the aircraft engine surges, the step-by-step increase in fuel supply is stopped and the aircraft engine surge relief plan is executed.

[0043] In some optional embodiments, in the above-mentioned method for obtaining the compressor stability boundary under the conditions of the entire aircraft engine, the flow rate and pressure ratio between the corresponding points on the working line and the points on the stability boundary at each given compressor conversion speed are obtained by interpolating the characteristics of the compressor components.

[0044] The pressure ratio of the compressor's stable operating point and surge operating point can be obtained based on the pressure parameters of the engine's routine test, and the compressor component characteristics can be obtained based on the compressor component test. The flow rates of the stable operating point and surge operating point are obtained by interpolation in the characteristic line diagram through the measured pressure ratio and speed. Then, the stable operating boundary of the compressor under various speed conditions can be obtained using the parameters obtained above.

[0045] The various embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to in detail.

[0046] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or replacements to the relevant technical features, and the technical solutions after these changes or replacements will fall within the scope of protection of the present application.

Claims

1. A method for obtaining the compressor stability boundary under the condition of an aircraft engine as a whole, characterized by: include: Make the aircraft engine run stably at a given compressor conversion speed; The fuel supply of the aircraft engine is increased stepwise to cause the aircraft engine to surge, and the stability boundary at a given compressor conversion speed is obtained; Change the given compressed air conversion speed and repeat the above steps to obtain the corresponding stability boundary; The fuel supply of the aircraft engine is increased in a stepwise manner to cause the aircraft engine to surge, and the stability boundary at a given compressor converted speed is obtained, specifically: Based on the stability margin designed for a given compressor converted speed line, the fuel supply rate of the aircraft engine is increased in a stepwise manner. The fuel supply rate is initially set based on the stability margin designed for the given compressor converted speed. When the aircraft engine does not experience surge, the value of the stepwise increase in fuel supply rate is increased until the aircraft engine does not experience surge, thereby obtaining the stability margin for the given compressor converted speed. The period of step-by-step increase in the fuel supply of an aircraft engine is in, ΔWFJY is the value of the step-wise increase in fuel supply; tHM is the actuation rate of the fuel supply valve mechanism of the aircraft engine control system; At each given compressor conversion speed, the flow rate and pressure ratio between the corresponding points on the working line and the points on the stability boundary are obtained by interpolating the characteristics of the compressor components.

2. The method for obtaining the compressor stability boundary under the condition of the entire aircraft engine according to claim 1, characterized in that: After the aircraft engine surges, the step-by-step increase in fuel supply is stopped and the aircraft engine surge relief plan is executed.

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