Airworthiness compliance verification method for surge and stall characteristics of aircraft engines during climb phase

Abstract

The present invention provides a method for verifying the airworthiness compliance of the surge and stall characteristics of an aircraft engine during the climb phase, comprising the following steps: S1, determining a verification engine; S2, simulating the aerodynamic flow field conditions of the verification engine during climb; S3, designing a test distortion generator; S4, using the test distortion generator to conduct a bench distortion test to calculate the surge margin of the test engine; S5, verifying whether the surge margin obtained by the test engine through step S4 meets the requirements. The present invention adopts a method combining analysis / calculation and (ground) distortion testing to verify the compliance of the engine's stability with the surge / stall clauses. Based on the evaluation of the engine performance during the design phase using distortion testing, combined with the minimum safety standards of the airworthiness clauses, its compliance with the airworthiness regulations is demonstrated, thereby also providing guarantees for the subsequent air trials, flight tests, and passing the final type certification.

Description

Technical Field

[0001] The invention belongs to the field of aircraft engine production, and in particular relates to a method for verifying the airworthiness compliance of surge and stall characteristics of aircraft engines during a climbing phase. Background Art

[0002] Part 33 of the Civil Aviation Regulations of China, CCAR-33-R2, "Aircraft Engine Airworthiness Regulations", as the minimum safety standard for aircraft engine airworthiness certification, sets minimum requirements for engine safety. In order for engine manufacturers to obtain a type certificate approved by the authority, they need to carry out relevant verification activities to demonstrate the compliance of aircraft engines with airworthiness clauses. However, what channels and methods are used to complete the verification of aircraft engine airworthiness clause compliance, what mechanisms are used to effectively track and manage safety demand data, verification data, process data, etc., and improve the authority's trust in compliance data, which are both key technologies that need to be solved urgently and difficult technologies.

[0003] Transport aircraft are mostly in a cruising state with stable working conditions, and the engine workload and inlet disturbance are very small. However, during the full-power climb process, the operating speed is the highest, the thrust is the highest, that is, the corresponding workload is the highest. At the same time, the inlet flow angle of the engine inlet increases, which may cause airflow separation, which will affect the surge / stall characteristics of the fan or compressor and the safety of the whole machine. Stall and surge are two typical unstable flow states in fans or compressors. Stall will cause fatigue of the blades under high-frequency alternating aerodynamic loads, which may cause blade damage. Surge will induce large fluctuations in engine pressure and flow, accompanied by periodic airflow oscillations, and the engine aerodynamic performance will be significantly deteriorated, and it may even cause the engine to shut down. Therefore, it is of positive significance to study the internal flow characteristics of the engine inlet and compression system during the aircraft climb process, and then develop the airworthiness compliance verification method of the engine under this condition. Summary of the invention

[0004] In view of this, the present invention aims to propose an airworthiness compliance verification method for the surge and stall characteristics of an aircraft engine during the aircraft climbing phase, so as to meet the compliance verification requirements for the airworthiness clauses of the aircraft engine.

[0005] To achieve the above object, the technical solution of the present invention is achieved as follows:

[0006] The method for verifying the airworthiness compliance of the surge and stall characteristics of aircraft engines during the climb phase includes the following steps:

[0007] S1. Determine the verification engine based on the test engine model and structural dimensions;

[0008] S2. Use numerical methods to simulate and verify the aerodynamic flow field conditions at different Mach numbers and different flight angles of attack under full engine power, and calculate the full-channel flow field under the conditions of internal and external flow coupling of the engine;

[0009] S3, determining a test distortion generator, using which the aerodynamic flow field condition described in step S2 can be simulated;

[0010] S4, applying the test distortion generator determined in step S3 to perform a bench distortion test, and calculating the surge margin of the test engine at full power under the aerodynamic flow field described in step S2 formed by the test distortion generator;

[0011] S5. According to the requirements of the airworthiness safety management system regarding surge / stall of aircraft engines during flight, verify whether the surge margin obtained in step S4 meets the requirements.

[0012] Furthermore, the method for determining the test distortion generator in step S3 includes the following steps:

[0013] S301, according to the result of step S2, obtaining an intake distortion map before verifying the engine fan under different Mach numbers and flight angles of attack, and obtaining a total pressure recovery coefficient and an intake distortion index of the map as design parameters;

[0014] S302, according to the intake distortion spectrum and design parameters obtained in step S301, designing a plurality of different distortion generators corresponding to each flight angle of attack condition at a corresponding Mach number;

[0015] S303, installing the distortion generators designed in step S302 one by one in front of the fan of the test engine in the test intake section, and calculating the intake distortion spectrum of the engine at full power state under different distortion generator schemes by studying different distortion generator and fan combination schemes, and obtaining the total pressure recovery coefficient of the spectrum section, the intake distortion index, and the pressure ratio and efficiency parameters of the fan;

[0016] S304, comparing the intake distortion spectra and related data of step S303 and step S301, and determining, among multiple distortion generators under corresponding Mach numbers and corresponding flight angles of attack, the distortion generator with the smallest parameter deviation from the one obtained in step S301 as the test distortion generator under the Mach number and the flight angle of attack;

[0017] Furthermore, the distortion generator is a plug-in type distortion generator.

[0018] Furthermore, the verification engine is a high bypass ratio turbofan engine with a nacelle.

[0019] Furthermore, the verification engine model is a DGEN380 aircraft engine.

[0020] Compared with the prior art, the method for verifying the airworthiness compliance of the surge and stall characteristics of an aircraft engine during the climb phase of the present invention has the following advantages:

[0021] In this design, numerical calculations are combined with bench distortion tests in the engine design and development process, and the airworthiness compliance verification of aircraft engines is integrated into the design and development process. Numerical calculations are used to obtain distortion maps under different flight Mach numbers and angles of attack and to design distortion generators. The surge margin of the engine under corresponding conditions is then determined through a distortion test program, thereby indicating compliance with the engine surge / stall clauses. Combined with the "Aircraft Type Certification Procedure" (AP-21-AA-2011-03-R4), a surge / stall compliance verification process for engines or compression components is developed, and data support and safety assurance are provided for subsequent air trials, flight tests and final type certification. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] The accompanying drawings constituting a part of the present invention are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the accompanying drawings:

[0023] Figure 1 The curves of the total pressure recovery coefficient σ and the inlet flow field distortion index D with the flight angle of attack AOA under different Mach numbers Ma of the engine DGEN380;

[0024] Figure 2 is the fan efficiency η of the DGEN380 engine at different angles of attack AOA when the Mach number is 0.4 * and pressure ratio π * characteristic line;

[0025] Figure 3 Distortion generators with two different board structures are designed;

[0026] Figure 4 It is the surge margin curve of the fan under different angles of attack and distortion test schemes;

[0027] Figure 5 It is a flow chart for the compliance verification of aircraft engine surge and stall;

[0028] Figure 6 This is a diagram of the bench distortion test device; DETAILED DESCRIPTION

[0029] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments may be combined with each other.

[0030] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside" and the like indicate positions or positional relationships based on the positions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the present invention. In addition, the terms "first", "second", and the like are only used for descriptive purposes, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, features defined as "first", "second", and the like may explicitly or implicitly include one or more of the features. In the description of the present invention, unless otherwise specified, "multiple" means two or more.

[0031] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it 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 it can be indirectly connected through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood by specific circumstances.

[0032] The present invention will be described in detail below with reference to the accompanying drawings and in conjunction with embodiments.

[0033] The airworthiness compliance verification method for the surge and stall characteristics of aircraft engines during the climb phase is as follows: Figure 5 As shown. Note: Before implementing this conformity verification method, a conformity plan must be formulated to determine the conformity verification object, as well as the clause requirements and clause intent that need to be met. Since the airworthiness certification, conformity verification methods and technologies of civil aviation engines are currently the weak areas in China, this conformity verification method is mainly aimed at civil high bypass ratio turbofan engines with nacelles, and verifies their compliance with surge and stall airworthiness clauses under the climbing condition of civil aircraft. Therefore, civil high bypass ratio turbofan engines with nacelles are selected as the conformity test objects. After the conformity plan is formulated, the specific implementation stage of the conformity method is entered, that is, the specific implementation stage of the airworthiness conformity verification method for surge and stall characteristics of aircraft engines in the climbing stage described in this embodiment is entered, which specifically includes the following steps:

[0034] S1. As the process Figure 5 As shown, determine the verification engine model based on the test engine model and structural dimensions, and obtain the verification engine overall or component characteristic requirements;

[0035] In this embodiment, the engine DGEN380 is selected as the conformity verification object, that is, the engine DGEN380 is the verification engine, and the design parameters and characteristics of the verification engine are obtained, for example: the fan design pressure ratio is 1.19, the design speed is 13069r / min, and the bypass ratio is 6.851, thereby completing the process Figure 5 The engine model determination link shown;

[0036] According to the process Figure 5 As shown in the requirements, after the test engine model is determined, the flight attitude of the aircraft in the flight phase and the working state of the verification engine must also be determined, including the flight altitude, Mach number, flight angle of attack and the working speed of the engine. In this embodiment, 100% of the design speed is selected, and the verification engine is in full power state at this time;

[0037] S2. Use numerical methods (such as three-dimensional numerical methods) to calculate and analyze the full-channel flow field under the conditions of internal and external flow coupling of the verification engine, obtain the aerodynamic characteristics and flow field characteristics of the verification engine during climb, and simulate the aerodynamic flow field conditions of different Mach numbers and different flight angles of attack under the full power state of the verification engine; thus completing the process Figure 5 The process characteristics and characteristic links of the engine are obtained through calculation and analysis;

[0038] S3, determining a test distortion generator, and simulating the aerodynamic flow field condition described in step S2 under the action of the test distortion generator;

[0039] The method for determining the test distortion generator includes the following steps:

[0040] S301, obtaining the intake distortion spectrum before the verification engine fan under different Mach numbers and flight angles of attack through the flow field characteristic results of step S2, and obtaining the change curves of the total pressure recovery coefficient σ and the intake distortion index D of the fan under different Mach numbers and different flight angles of attack with the flight angle of attack AOA, such as Figure 1 As shown, the total pressure recovery coefficient σ and the intake distortion index D of the spectrum are obtained as design parameters; for the convenience of explanation, this embodiment takes the Mach number of 0.4 as an example, and calculates the efficiency η of the fan under various flight angles of attack AOA conditions when the Mach number is 0.4. * and pressure ratio π * Characteristic curve, see Figure 2 As shown;

[0041] S302: Based on the intake distortion map and design parameters obtained in step S301, a plurality of different distortion generators are designed for each flight angle of attack. This step is the process Figure 5In the link of designing a distortion generator according to the distortion spectrum; in this link, the multiple different distortion generators corresponding to the same flight angle of attack may be distortion generators with different structures, or may be distortion generators with the same structure but different specific sizes, that is, the distortion generator may be designed in different structural forms, or the distortion generator of the same structure may be designed in multiple sizes, that is, for the distortion generator of the same structure, the specific size parameters are different; in this embodiment, the distortion generator is a plug-in type distortion generator, such as Figure 3 Shown are two types of distortion generators with different plug-in board structures;

[0042] S303, installing the distortion generators designed in step S302 one by one in front of the fan of the test engine in the test intake section, and calculating the intake distortion spectrum of the engine at full power under different distortion generator schemes to obtain the total pressure recovery coefficient σ, the intake distortion index D, and the pressure ratio and efficiency parameters of the fan in the spectrum section;

[0043] S304, comparing the intake distortion spectra and related data of step S303 and step S301, and determining the distortion generator with the smallest parameter deviation from the obtained in step S301 among multiple different distortion generators at each flight angle of attack when the Mach number is 0.4 as the test distortion generator at the Mach number of 0.4 and the flight angle of attack, and determining the test distortion generator at each flight angle of attack when the Mach number is 0.4 by analogy;

[0044] The implementation process of steps S303 and S304 is the process Figure 5 The feasibility verification link of the distortion generator is shown in , in which the distortion generator that passes the feasibility verification is the test distortion generator;

[0045] In step S303, by adjusting the installation position of each distortion generator (the circumferential position of the distortion generator in the test air intake section and the axial distance of the distortion generator from the installation position of the test engine fan), it is ensured that the aerodynamic flow field simulated at this position under the condition of the distortion generator is closest to the aerodynamic flow field under the same Mach number and the same flight angle of attack of the verification engine, and the distortion map parameters and fan characteristic curve (such as the total pressure recovery coefficient σ, the intake distortion index D, and the pressure ratio and efficiency parameters of the fan) under the distortion generator and their installation position information are recorded. After the test of a distortion generator is completed, it is taken out and replaced with other distortion generators with the same corresponding flight angle of attack for testing, and so on until the test of all distortion generators under the flight angle of attack condition is completed, and the flow field parameters of all distortion generators under the flight angle of attack condition are recorded, and the position information of all distortion generators under the flight angle of attack condition is recorded; in step S304, the flow field parameters of each distortion generator under the same flight angle of attack obtained in step S303 are compared with the relevant parameters in the intake distortion spectrum obtained in step S301, and finally the distortion generator with the smallest difference in the comparison results is selected as the test distortion generator under the flight angle of attack condition; in this way, the test distortion generators under different flight angles of attack conditions are determined, and the installation position information of each test distortion generator is recorded, which is the test installation position of the test distortion generator when the bench distortion test is carried out;

[0046] S4, using the test distortion generator determined in step S3 to perform a bench distortion test, this step is the process Figure 5 In the process of using the distortion generator to conduct bench distortion test and obtain surge margin characteristics, the specific operation method is: install the test distortion generator to the test installation position of the test intake section, such as Figure 6 As shown in FIG. 1 , the test engine is in full power state, and the surge margin SM value of the test engine during full power climbing under the aerodynamic flow field formed by different test distortion generators is obtained, and the surge margin SM curve is drawn, as shown in FIG. Figure 4 As shown in the figure, the surge margin value under the condition of extreme intake distortion, i.e. when the climb angle of attack is maximum, is further determined; Figure 4 The other curve is the margin curve of engine DGEN380 at different angles of attack. It can be seen that under the same flight angle of attack conditions, the margin value of the test engine is close to the margin value of engine DGEN380, with a maximum deviation of 4.32%, and the change trends of the two curves are basically the same.

[0047] S5, as the process Figure 5The last two steps shown are to verify whether the surge margin SM obtained in step S4 meets the requirements of the provisions on aircraft engine surge / stall in the airworthiness management system (such as the Civil Aviation Regulations of China) (i.e., the provisions of CCAR-33-R2 "Aircraft Engine Airworthiness Regulations"), and submit a compliance verification report and conclusion; if it meets the requirements, the test engine can be subjected to subsequent tests, such as air tests, flight tests, etc., and provide data support and safety assurance for the final type certification; if it does not meet the requirements, it means that the engine safety does not meet the standards.

[0048] In this embodiment, the engine intake duct total pressure recovery coefficient is defined as follows:

[0049]

[0050] in is the average total pressure at the inlet outlet section, Indicates the total free flow pressure far ahead.

[0051] The intake distortion index is defined as follows:

[0052]

[0053] in and They represent the maximum and minimum total pressures at the inlet duct outlet section respectively.

[0054] Surge margin refers to the difference between the pressure ratio on the stable boundary line and the common working line divided by the pressure ratio on the common working line. The definition formula is as follows based on the same conversion speed and considering the influence of the conversion flow change:

[0055]

[0056] In the formula and are the near-stall flow rate and total pressure ratio, and is the design operating flow rate and total pressure ratio.

[0057] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. The method for verifying the airworthiness compliance of surge and stall characteristics of aircraft engines during the climb phase is characterized by: The following steps are involved: S1. Determine the verification engine based on the test engine model and structural dimensions; S2. Use numerical methods to simulate and verify the aerodynamic flow field conditions at different Mach numbers and different flight angles of attack under full engine power, and calculate the full-channel flow field under the conditions of internal and external flow coupling of the engine; S3, determining a test distortion generator, using which the aerodynamic flow field condition described in step S2 can be simulated; S4, applying the test distortion generator determined in step S3 to perform a bench distortion test on the test engine, and calculating the surge margin of the test engine at full power under the aerodynamic flow field in step S2 formed by the test distortion generator; S5. According to the requirements of the airworthiness safety management system regarding surge / stall of aircraft engines during flight, verify whether the surge margin of the test engine obtained in step S4 meets the requirements; The method for determining the test distortion generator in step S3 includes the following steps: S301, according to the result of step S2, obtaining an intake distortion map before verifying the engine fan under different Mach numbers and flight angles of attack, and obtaining a total pressure recovery coefficient and an intake distortion index of the map as design parameters; S302, according to the intake distortion map and design parameters obtained in step S301, designing a plurality of different distortion generators corresponding to different flight angles of attack under corresponding Mach numbers; S303, installing the distortion generators designed in step S302 one by one in front of the fan of the test engine in the test intake section, and calculating the intake distortion spectrum of the test engine at full power state under different distortion generator schemes by studying different distortion generator and fan combination schemes, and obtaining the total pressure recovery coefficient of the spectrum section, the intake distortion index, and the pressure ratio and efficiency parameters of the fan; S304, comparing the intake distortion spectra and related data of step S303 and step S301, and determining, among multiple distortion generators under corresponding Mach numbers and corresponding flight angles of attack, the distortion generator with the smallest parameter deviation from that obtained in step S301 as the test distortion generator under the Mach number and the flight angle of attack.

2. The method for verifying the airworthiness compliance of surge and stall characteristics of an aircraft engine during the climb phase according to claim 1, characterized in that: The distortion generator is a plug-in type distortion generator.

3. The method for verifying the airworthiness compliance of surge and stall characteristics of an aircraft engine during the climb phase according to claim 1, characterized in that: The verification engine is a high bypass ratio turbofan engine with a nacelle.

4. The method for verifying the airworthiness compliance of surge and stall characteristics of an aircraft engine during the climb phase according to claim 3, characterized in that: The verification engine model is the DGEN380 aircraft engine.

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