A method for derivative development design of a lower pressure flow path of an axial flow compressor of an aero-engine

Abstract

The application relates to a design method for deriving and developing a downward flow path of an aero-engine axial flow compressor, which comprises the following steps: step one, downward pressing of compressor inlet and outlet hubs, namely, downward pressing of inner flow paths of the compressor inlet and outlet; step two, re-determination of the compressor hub flow path according to the downward pressing amount of the compressor inlet and outlet hubs; step three, re-determination of the compressor casing flow path while keeping the compressor flow path area unchanged; step four, re-determination of the design rotating speed of the compressor while keeping the compressor rotating speed inlet blade tip tangential velocity unchanged; step five, adjustment of the blade profile of each stage of rotor and stator blades while keeping the compressor stage pressure ratio, efficiency and radial distribution unchanged; step six, adjustment of the inlet pre-whirl angle of each stage while keeping the compressor rotor counterforce unchanged; and step seven, checking of the radial distribution of the diffusion factors of each stage of the compressor rotor and stator, and if the radial distribution exceeds the engineering applicable value, returning to step one, adjusting the downward pressing amount of the compressor inlet and outlet hubs and performing iterative optimization.

Description

Technical Field

[0001] This application belongs to the field of aero-engine axial compressor design technology, specifically relating to a design method for the derivation and development of the downflow path of an aero-engine axial compressor. Background Technology

[0002] Axial compressors for aero engines have multiple stages and high performance requirements. Designing a completely new multi-stage axial compressor requires significant investment, has a long design cycle, and carries high design risks. Developing derivatives based on existing compressors is a convenient way to achieve rapid compressor design.

[0003] The main methods for developing derivatives based on existing compressors are geometric scaling and compressor tip cutting. Among them, geometric scaling follows the principle of compressor geometric similarity and has a good theoretical foundation; compressor tip cutting involves cutting off the top of the compressor, removing the tip of the existing compressor, which is simple to operate.

[0004] Currently, for axial compressors of aero engines, the goal is often to design them with low weight while ensuring flow rate and aerodynamic performance. However, the compressor flow rate changes with the square of the geometric scaling ratio. Removing the compressor tip will reduce the compressor flow rate. Geometric scaling and compressor tip cutting methods cannot design compressors with low flow resistance and weight while ensuring flow rate and aerodynamic performance.

[0005] In view of the above, this application is hereby filed. Summary of the Invention

[0006] The purpose of this application is to provide a design method for the derivation and development of the pressure flow path of an axial compressor for an aero-engine. Based on the existing compressor, a compressor with lower weight can be designed while ensuring the flow rate and its aerodynamic performance, so as to quickly and conveniently realize the weight reduction design of the compressor.

[0007] The technical solution of this application is:

[0008] A design method for the development of a downflow path in an axial compressor of an aero-engine includes:

[0009] Step 1: Press down the compressor inlet and outlet hubs, which means press down the internal flow path of the compressor inlet and outlet.

[0010] Step 2: Based on the downward pressure at the compressor inlet and outlet hubs, redetermine the compressor hub flow path;

[0011] Step 3: Keep the compressor flow path area unchanged and redetermine the compressor casing flow path;

[0012] Step 4: Keep the compressor speed and inlet blade tip tangential velocity constant, and redetermine the compressor design speed.

[0013] According to at least one embodiment of this application, the above-described aero-engine axial compressor further includes:

[0014] Step 5: Adjust the blade profile of each stage of the compressor stator to keep the pressure ratio, efficiency and radial distribution of each stage of the compressor constant;

[0015] Step 6: Adjust the inlet pre-swirl angle of each stage to keep the rotor counterforce of each stage of the compressor constant.

[0016] According to at least one embodiment of this application, the above-described aero-engine axial compressor further includes:

[0017] Step 7: Check the radial distribution of the diffusion factors of the rotor and stator at each stage of the compressor. If it exceeds the applicable engineering value, return to Step 1 and adjust the downward pressure on the compressor inlet and outlet hubs for iterative optimization.

[0018] According to at least one embodiment of this application, in the first step of the above-mentioned aero-engine axial compressor, the downward pressure H1 of the compressor inlet hub is less than 20% of the original compressor inlet hub radius r11, the downward pressure H2 of the compressor outlet hub is less than 20% of the original compressor outlet hub radius r12, and H1≥H2.

[0019] According to at least one embodiment of this application, in step one of the above-described aero-engine axial compressor:

[0020] H2 = 0.8 * H1, or .

[0021] According to at least one embodiment of this application, in step three of the above-described aero-engine axial compressor process, for any axial position of the compressor:

[0022] ;

[0023] Where R2new is the radius of the compressor casing flow path that has been redefined; R1new is the radius of the compressor hub flow path that has been redefined; R2 is the radius of the original compressor casing flow path; and R1 is the radius of the original compressor hub flow path.

[0024] According to at least one embodiment of this application, in the above-described method for the derivation and development of the downflow path of an aero-engine axial compressor, step four includes:

[0025] ;

[0026] Where Nnew is the redefined design speed of the compressor; Nold is the original design speed of the compressor; R21 is the original inlet casing radius of the compressor; and R21new is the redefined inlet casing radius of the compressor.

[0027] This application has at least the following beneficial technical effects:

[0028] This paper presents a design method for the derivation and development of the pressure flow path of an axial compressor for aero-engines. Based on existing compressors, it can quickly build a new compressor flow field design scheme with the same flow rate and pressure ratio while ensuring the flow rate and aerodynamic performance. This allows for the design of a compressor with lower weight, achieving compressor weight reduction design. Furthermore, it can fully inherit the design parameters of existing compressors, and the performance of the redesigned compressor can be basically equivalent to the original scheme. This can greatly accelerate the compressor design cycle and reduce design risks. Attached Figure Description

[0029] Figure 1 This is a schematic diagram of the design method for the derivation and development of the downflow path of the axial compressor of an aero-engine provided in the embodiments of this application.

[0030] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0031] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0032] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0033] A design method for the development of the downflow path in an axial compressor of an aero-engine, such as... Figure 1 As shown.

[0034] Step 1: Press down the compressor inlet and outlet hubs, which means press down the internal flow path of the compressor inlet and outlet.

[0035] The downward pressure H1 at the compressor inlet hub is less than 20% of the original compressor inlet hub radius r11, and the downward pressure H2 at the compressor outlet hub is less than 20% of the original compressor outlet hub radius r12, and H1 ≥ H2. Typically, H2 = 0.8 * H1, or... .

[0036] Step 2: Based on the downward pressure of the compressor inlet and outlet hubs, redetermine the compressor hub flow path, which is to redetermine the internal flow path of the compressor.

[0037] Based on the downward pressure at the compressor inlet and outlet hubs, linear interpolation is used to obtain the downward pressure of the hub flow path at each axial position, thereby redetermining the compressor hub flow path.

[0038] Step 3: Keeping the compressor flow path area constant, redetermine the compressor casing flow path, i.e., redetermine the compressor external flow path. For any axial position:

[0039] ;

[0040] Where R2new is the radius of the compressor casing flow path that has been redefined; R1new is the radius of the compressor hub flow path that has been redefined; R2 is the radius of the original compressor casing flow path; and R1 is the radius of the original compressor hub flow path.

[0041] Step 4: Keeping the compressor speed and inlet blade tip tangential velocity constant, redetermine the compressor's design speed, which includes:

[0042] ;

[0043] Where Nnew is the redefined design speed of the compressor; Nold is the original design speed of the compressor; R21 is the original inlet casing radius of the compressor; and R21new is the redefined inlet casing radius of the compressor.

[0044] Step 5: Adjust the blade profile of each stage of the compressor stator to keep the pressure ratio, efficiency and radial distribution of each stage of the compressor unchanged.

[0045] Step 6: Adjust the inlet pre-swirl angle of each stage to keep the rotor counterforce of each stage of the compressor constant.

[0046] The input parameters for compressor design at each stage, including the number of blades, blade angle of attack, and mid-curvature, can be kept consistent.

[0047] Step 7: The design of the down-pressure flow path will increase the aerodynamic load of the compressor. Check the radial distribution of the diffusion factor of the rotor and stator of each stage of the compressor. If it exceeds the applicable value for engineering, return to step 1 and adjust the down-pressure amount of the compressor inlet and outlet hubs for iterative optimization.

[0048] The design method for the derivation and development of the pressure flow path of the axial compressor of an aero-engine disclosed in the above embodiments can, based on the existing compressor, quickly build a new compressor flow field design scheme with the same flow rate and pressure ratio while ensuring the flow rate and its aerodynamic performance. This allows for the design of a compressor with a lower weight, achieving compressor weight reduction design. Furthermore, it can fully inherit the design parameters of the existing compressor, and the performance of the redesigned compressor can be basically equivalent to the original scheme. This can greatly accelerate the compressor design cycle and reduce design risks.

[0049] The technical solution of this 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 this application is obviously not limited to these specific embodiments. Without departing from the principles of this 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 all fall within the scope of protection of this application.

Claims

1. A design method for the derivation and development of the downflow path of an axial compressor in an aero-engine, characterized in that, include: Step 1: Press down the compressor inlet and outlet hubs, which means press down the internal flow path of the compressor inlet and outlet. Step 2: Based on the downward pressure at the compressor inlet and outlet hubs, redetermine the compressor hub flow path; Step 3: Keep the compressor flow path area unchanged and redetermine the compressor casing flow path; Step 4: Keep the compressor speed and inlet blade tip tangential velocity constant, and redetermine the compressor design speed; Step 5: Adjust the blade profile of each stage of the compressor stator to keep the pressure ratio, efficiency and radial distribution of each stage of the compressor constant; Step 6: Adjust the inlet pre-swirl angle of each stage to maintain the constant counterforce of the compressor rotor at each stage; Step 7: Check the radial distribution of the diffusion factors of the rotor and stator at each stage of the compressor. If it exceeds the applicable engineering value, return to Step 1 and adjust the downward pressure on the compressor inlet and outlet hubs for iterative optimization.

2. The design method for the derivation and development of the downflow path of an aero-engine axial compressor according to claim 1, characterized in that, In step one, the downward pressure H1 of the compressor inlet hub is less than 20% of the original compressor inlet hub radius r11, and the downward pressure H2 of the compressor outlet hub is less than 20% of the original compressor outlet hub radius r12, and H1≥H2.

3. The design method for the derivation and development of the downflow path of an aero-engine axial compressor according to claim 2, characterized in that, In step one: H2 = 0.8 * H1, or .

4. The design method for the derivation and development of the downflow path of an aero-engine axial compressor according to claim 3, characterized in that, In step three, for any axial position of the compressor: ； Where R2new is the radius of the compressor casing flow path that has been redefined; R1new is the radius of the compressor hub flow path that has been redefined; R2 is the radius of the original compressor casing flow path; and R1 is the radius of the original compressor hub flow path.

5. The design method for the derivation and development of the downflow path of an aero-engine axial compressor according to claim 4, characterized in that, Step four includes: ； Where Nnew is the redefined design speed of the compressor; Nold is the original design speed of the compressor; R21 is the original inlet casing radius of the compressor; and R21new is the redefined inlet casing radius of the compressor.

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