A method for controlling the angle of adjustable blades of an aircraft engine compressor

Abstract

The present application belongs to the technical field of adjustable blade angle control of an aircraft engine compressor, and specifically relates to a method for controlling the adjustable blade angle of an aircraft engine compressor, comprising: performing an aircraft engine whole machine test to obtain a first compressor adjustable blade angle control law a 2_1 =f(p 3_av / p1), where p 3_av is the mean value of the compressor outlet pressure distribution; p1 is the aircraft engine inlet pressure; the aircraft engine whole machine test is used to obtain the second compressor adjustable blade angle control law a 2_2 =f(N 2R_k ), N 2R_k =N2·(288.15 / T 1_max ) 0.5 , where N2 is the high pressure speed of the aircraft engine; T 1_max is the maximum value of the aero-engine inlet temperature distribution; through the aero-engine component test, the control law a of the adjustable blade angle of the third compressor is obtained 2_3 =f(N 2R25_k ), N 2R25_k =k·N2·(288.15 / T 25 ) 0.5 , where k is the aircraft engine speed coefficient; T 25 is the compressor inlet temperature; a 2_1 、a 2_2 、a 2_3 The larger value is used as the control angle a2 of the compressor adjustable blades.

Description

Technical Field

[0001] The present application belongs to the technical field of adjustable blade angle control of an aircraft engine compressor, and specifically relates to a method for controlling the adjustable blade angle of an aircraft engine compressor. Background Art

[0002] The aircraft engines of carrier-based aircraft are equipped with deflector baffles, which are used to block the airflow ejected from the tail nozzle of the aircraft engine when the carrier-based aircraft takes off from the deck, so as to avoid damage to the deck equipment and nearby personnel. However, this part of the airflow may flow to the aircraft engine inlet, causing the aircraft engine intake temperature to rise significantly, resulting in distortion of the aircraft engine inlet temperature.

[0003] Currently, most of them are based on the aircraft engine inlet temperature T 1 Or compressor inlet temperature T 25 The control law of the compressor adjustable blade angle is obtained to control the angle of the compressor adjustable blade in the aircraft engine. When the aircraft takes off from the deck of the carrier aircraft and the aircraft engine inlet temperature is distorted, the aircraft engine inlet temperature T 1 , may be lower than the average temperature of the aircraft engine inlet section. In this case, based on the aircraft engine inlet temperature T 1 The angle of the compressor adjustable blade is obtained. Controlling the angle of the compressor adjustable blade in the aircraft engine will make the compressor adjustable blade angle relatively biased, resulting in a large compressor margin loss. Based on the compressor inlet temperature T 25 The control law of the compressor is used to obtain the angle of the compressor adjustable blades. The angle of the compressor adjustable blades in the aircraft engine is controlled. The compressor inlet temperature T 25 The time response constant is large, and there is also the problem of causing the angle of the compressor adjustable blades to be relatively deviated, resulting in a large compressor margin loss.

[0004] This application is proposed in view of the above-mentioned technical defects.

[0005] 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 the present application. Summary of the invention

[0006] The purpose of the present application is to provide a method for controlling the angle of adjustable blades of an aircraft engine compressor, so as to overcome or alleviate at least one of the known technical defects.

[0007] The technical solution of this application is:

[0008] A method for controlling the angle of adjustable blades of an aircraft engine compressor, comprising:

[0009] The control law of the adjustable blade angle of the first compressor is obtained by the whole-machine test of the aircraft engine. 2_1 =f(p 3_av / p 1 ), where a 2_1 is the compressor adjustable blade angle obtained according to the first compressor adjustable blade angle control law; p 3_av is the mean value of the compressor outlet pressure distribution; p 1 is the aircraft engine inlet pressure;

[0010] The control law of the adjustable blade angle of the second compressor is obtained by the whole-machine test of the aircraft engine. 2_2 =f(N 2R_k ), N 2R_k =N 2 ·(288.15 / T 1_max ) 0.5 , where a 2_2 is the compressor adjustable blade angle obtained according to the second compressor adjustable blade angle control law; N 2 is the high pressure speed of the aircraft engine; T 1_max is the maximum value of the aircraft engine inlet temperature distribution;

[0011] The control law of the adjustable blade angle of the third compressor is obtained by testing the aircraft engine components. 2_3 =f(N 2R25_k ), N 2R25_k = k·N 2 ·(288.15 / T 25 ) 0.5 , where a 2_3 The adjustable compressor blade angle is obtained according to the third compressor adjustable blade angle control law; k is the aircraft engine speed coefficient; T 25 is the compressor inlet temperature;

[0012] With a 2_1 、a 2_2 、a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2 .

[0013] According to at least one embodiment of the present application, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, the mean value p of the compressor outlet pressure distribution is 3_av , obtained by averaging the pressures measured by pressure sensors arranged circumferentially at the compressor outlet.

[0014] According to at least one embodiment of the present application, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, the maximum value T of the aircraft engine inlet temperature distribution is 1_max , the value is the maximum value measured by the temperature sensors arranged circumferentially at the inlet of the aircraft engine.

[0015] According to at least one embodiment of the present application, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, the value of the aircraft engine speed coefficient k is between 0.5 and 1.5.

[0016] According to at least one embodiment of the present application, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, N is obtained by testing aircraft engine components. 1R_K =f(T 25 / T 1_max ), N 1R_K =N 1 ·(288.15 / T 1_max ) 0.5 , where N 1 It is the low pressure speed of the aircraft engine;

[0017] Compressor inlet temperature T 25 By N 1R_K , T 1_max , with N 1R_K =f(T 25 / T 1_max ) is concluded.

[0018] This application has at least the following beneficial technical effects:

[0019] A method for controlling the angle of adjustable blades of an aircraft engine compressor is provided. The method is designed based on the mean value p of the compressor outlet pressure distribution. 3_av , aircraft engine inlet pressure p 1 The first compressor adjustable blade angle control law obtains the compressor adjustable blade angle a 2_1 , based on the aircraft engine high pressure speed N 2 , the maximum value of the aircraft engine inlet temperature distribution T 1_max The second compressor adjustable blade angle control law obtains the compressor adjustable blade angle a 2_2 , based on the aircraft engine high pressure speed N 2 , compressor inlet temperature T 25 The third compressor adjustable blade angle control law obtains the compressor adjustable blade angle a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2In this way, when a carrier-based aircraft takes off from the deck and the inlet temperature of the aircraft engine is distorted, the angle of the adjustable compressor blades can be controlled in a timely and effective manner, thereby reducing the compressor margin loss, improving the aerodynamic stability of the aircraft engine, and increasing the ability to resist temperature distortion. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a schematic diagram of an adjustable blade angle control method for an aircraft engine compressor provided in an embodiment of the present application.

[0021] 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 cannot be understood as limitations on this patent. DETAILED DESCRIPTION

[0022] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be further described in detail in detail and in detail in conjunction with the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience 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 in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0023] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall be the common meanings understood by the general technicians in the field to which this application belongs. The words "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inside", "outside" and other words indicating orientation used in the description of this application are only used to indicate the relative direction or positional relationship, and do not imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, so it cannot be understood as a limitation on this application. The "first", "second", "third" and similar terms used in the description of this application are only used for descriptive purposes to distinguish different components, and cannot be understood as indicating or implying relative importance. The similar words "one", "one" or "the" used in the description of this application should not be understood as an absolute limitation on quantity, but should be understood as the existence of at least one. The similar words "including" or "comprising" used in the description of this application mean that the elements or objects appearing in front of the word include the elements or objects listed after the word and their equivalents, without excluding other elements or objects.

[0024] In addition, it should be noted that, unless otherwise clearly specified and limited, the words "installed", "connected", "connected" 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 the internal connection of two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.

[0025] The following is combined with Figure 1 This application is described in further detail.

[0026] A method for controlling the angle of adjustable blades of an aircraft engine compressor, comprising:

[0027] The control law of the adjustable blade angle of the first compressor is obtained by the whole-machine test of the aircraft engine. 2_1 =f(p 3_av / p 1 ), where a 2_1 is the compressor adjustable blade angle obtained according to the first compressor adjustable blade angle control law; p 3_av is the mean value of the compressor outlet pressure distribution; p 1 is the aircraft engine inlet pressure;

[0028] The control law of the adjustable blade angle of the second compressor is obtained by the whole-machine test of the aircraft engine. 2_2 =f(N 2R_k ), N 2R_k =N 2 ·(288.15 / T 1_max ) 0.5 , where a 2_2 is the compressor adjustable blade angle obtained according to the second compressor adjustable blade angle control law; N 2 is the high pressure speed of the aircraft engine; T 1_max is the maximum value of the aircraft engine inlet temperature distribution;

[0029] Through the aircraft engine test, the control law of the adjustable blade angle of the third compressor is obtained. 2_3 =f(N 2R25_k ), N 2R25_k = k·N 2 ·(288.15 / T 25 ) 0.5 , where a 2_3 The adjustable compressor blade angle is obtained according to the third compressor adjustable blade angle control law; k is the aircraft engine speed coefficient; T 25 is the compressor inlet temperature;

[0030] With a2_1 、a 2_2 、a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2 .

[0031] For the method for controlling the adjustable blade angle of an aircraft engine compressor disclosed in the above embodiment, it can be understood by those skilled in the art that the mean value p of the compressor outlet pressure distribution is selected. 3_av , aircraft engine inlet pressure p 1 The first compressor adjustable blade angle control law obtains the compressor adjustable blade angle a 2_1 , based on the aircraft engine high pressure speed N 2 , the maximum value of the aircraft engine inlet temperature distribution T 1_max The second compressor adjustable blade angle control law obtains the compressor adjustable blade angle a 2_2 , based on the aircraft engine high pressure speed N 2 , compressor inlet temperature T 25 The third compressor adjustable blade angle control law obtains the compressor adjustable blade angle a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2 In this way, when a carrier-based aircraft takes off from the deck and the inlet temperature of the aircraft engine is distorted, the angle of the adjustable compressor blades can be controlled in a timely and effective manner, thereby reducing the compressor margin loss, improving the aerodynamic stability of the aircraft engine, and increasing the ability to resist temperature distortion.

[0032] In order to achieve further better results, 1_max >T 1_av +A, take a 2_1 、a 2_2 、a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2 , in T 1_max ≤T 1_av +A, take a 2_2 、a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2 ;in,

[0033] T 1_av is the mean value of the aircraft engine inlet temperature distribution;

[0034] A is a threshold value set to determine whether there is an increase in the aircraft engine inlet temperature, and the specific value may be 10°C to 30°C.

[0035] In some optional embodiments, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, the mean value p of the compressor outlet pressure distribution is3_av The pressure is measured by arranging pressure sensors along the circumference of the compressor outlet and taking the average value, which is easy to measure.

[0036] In some optional embodiments, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, the maximum value T of the aircraft engine inlet temperature distribution is 1_max , the value is taken as the maximum value measured by the temperature sensors arranged along the circumference of the aircraft engine inlet, so as to avoid the temperature of the aircraft engine inlet being lower than the average temperature of the aircraft engine inlet section when the aircraft engine inlet temperature is distorted.

[0037] In some optional embodiments, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, the value of the aircraft engine speed coefficient k is between 0.5 and 1.5.

[0038] In some optional embodiments, in the above-mentioned method for controlling the adjustable blade angle of an aircraft engine compressor, N is obtained by testing aircraft engine components. 1R_K =f(T 25 / T 1_max ), N 1R_K =N 1 ·(288.15 / T 1_max ) 0.5 , where N 1 It is the low pressure speed of the aircraft engine;

[0039] Compressor inlet temperature T 25 By N 1R_K , T 1_max , with N 1R_K =f(T 25 / T 1_max ) is obtained, which is convenient for obtaining and reducing the compressor inlet temperature T 25 The time response constant.

[0040] 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 referenced to each other.

[0041] 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 substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.

Claims

1. A method for controlling the angle of adjustable blades of an aircraft engine compressor. It is characterized in that include: The control law of the adjustable blade angle of the first compressor is obtained by the whole-machine test of the aircraft engine. 2_1 =f(p 3_av / p 1 ), where a 2_1 is the compressor adjustable blade angle obtained according to the first compressor adjustable blade angle control law; p 3_av is the mean value of the compressor outlet pressure distribution; p 1 is the aircraft engine inlet pressure; The control law of the adjustable blade angle of the second compressor is obtained by the whole-machine test of the aircraft engine. 2_2 =f(N 2R_k ), N 2R_k =N 2 ·(288.15 / T 1_max ) 0.5 , where a 2_2 is the compressor adjustable blade angle obtained according to the second compressor adjustable blade angle control law; N 2 is the high pressure speed of the aircraft engine; T 1_max is the maximum value of the aircraft engine inlet temperature distribution; The control law of the adjustable blade angle of the third compressor is obtained by testing the aircraft engine components. 2_3 =f(N 2R25_k ), N 2R25_k = k·N 2 ·(288.15 / T 25 ) 0.5 , where a 2_3 The adjustable compressor blade angle is obtained according to the third compressor adjustable blade angle control law; k is the aircraft engine speed coefficient; T 25 is the compressor inlet temperature; With a 2_1 、a 2_2 、a 2_3 The larger value is used as the control angle a of the compressor adjustable blade 2 .

2. The method for controlling the adjustable blade angle of an aircraft engine compressor according to claim 1, It is characterized in that The mean value p of the compressor outlet pressure distribution 3_av , obtained by averaging the pressures measured by pressure sensors arranged circumferentially at the compressor outlet.

3. The method for controlling the adjustable blade angle of an aircraft engine compressor according to claim 1, It is characterized in that Maximum value T of the aircraft engine inlet temperature distribution 1_max , the value is the maximum value measured by the temperature sensors arranged circumferentially at the inlet of the aircraft engine.

4. The method for controlling the adjustable blade angle of an aircraft engine compressor according to claim 1, It is characterized in that The value of the aircraft engine speed coefficient k is between 0.5 and 1.

5.

5. The method for controlling the adjustable blade angle of an aircraft engine compressor according to claim 1, It is characterized in that Using aerospace engine parts testing, we obtained N 1R_K =f(T 25 / T 1_max ), N 1R_K =N 1 ·(288.15 / T 1_max ) 0.5 , where N 1 It is the low pressure speed of the aircraft engine; Compressor inlet temperature T 25 By N 1R_K , T 1_max , with N 1R_K =f(T 25 / T 1_max ) is concluded.

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