A method for determining the ground intake air heating temperature during aircraft engine endurance test

Abstract

The present application belongs to the field of engine ground testing, and relates to a method for determining the ground intake heating temperature of an aircraft engine endurance test. The method comprises: step S1, determining the engine inlet temperature and engine inlet pressure corresponding to the flight mission; step S2, determining the compressor outlet gas parameters and the gas parameters before the main flow channel turbine rotor; step S3, calculating the cooling gas parameters introduced into the turbine rotor; step S4, calculating the theoretical turbine matrix temperature; step S5, determining the ground turbine matrix temperature calculated under the given ground normal pressure and initial heating temperature conditions during the aircraft engine endurance test; step S6, adjusting the initial heating temperature of the test so that the theoretical turbine matrix temperature is the same as the ground turbine matrix temperature. The present application ensures that the damage in the endurance test fully simulates the actual damage in the field, which can fully assess the engine durability without causing excessive assessment, and effectively ensure flight safety.

Description

Technical Field

[0001] The present application belongs to the field of engine ground testing, and in particular relates to a method for determining the ground intake heating temperature of an aircraft engine endurance test. Background Art

[0002] Aircraft engines are generally verified for structural durability through ground endurance tests, including low-cycle fatigue, creep / stress rupture, etc., to simulate the endurance creep damage (mainly hot parts damage) during actual use. Currently, the engine inlet air temperature is made the same as the corresponding flight mission state through intake heating tests, and the influence of inlet pressure is usually not considered.

[0003] At present, the intake heating temperature of the endurance test is selected according to the inlet temperature of the same air flight mission. Although the gas temperature of the engine flow path is the same as the corresponding flight mission state, the different turbine cooling air pressures and turbine cooling effects result in different temperatures of the turbine blade bases, which are different from the thermal damage in actual flight. If the damage is less than that in the air, the assessment will not be in place, and the actual durability problems will not be fully exposed, which cannot effectively guarantee flight safety. If the damage is greater than that in the air, it will lead to excessive assessment, which may lead to failure of the endurance test and affect the development progress. Summary of the invention

[0004] In order to solve the above problems, the present application provides a method for determining the ground intake heating temperature for a durable test of an aircraft engine, which solves the problem that the existing method for determining the intake heating temperature causes damage during the durable test and deviates from actual use.

[0005] The present application provides a method for determining the ground intake air heating temperature of an aircraft engine endurance test, which mainly includes:

[0006] Step S1, determining the engine inlet temperature and the engine inlet pressure at each design point corresponding to the flight mission simulated by the aircraft engine endurance test;

[0007] Step S2, determining the compressor outlet gas parameters and the main flow channel turbine rotor front gas parameters according to the engine inlet temperature and the engine inlet pressure;

[0008] Step S3, calculating the cooling air parameters introduced into the turbine rotor according to the compressor outlet parameters and the flow path parameters of the cooling system for cooling the turbine rotor through the compressor bleed air;

[0009] Step S4, calculating the theoretical turbine base temperature according to the cooling gas parameters and the gas parameters before the main flow channel turbine rotor;

[0010] Step S5, determining the ground turbine base temperature calculated under the given ground normal pressure and initial heating temperature conditions during the aircraft engine endurance test;

[0011] Step S6: For each design point, the initial heating temperature of the test is adjusted according to the theoretical turbine matrix temperature and the ground turbine matrix temperature, so that the theoretical turbine matrix temperature is the same as the ground turbine matrix temperature, thereby forming the initial heating temperature of each design point of the aircraft engine endurance test.

[0012] Preferably, step S1 further comprises:

[0013] Step S11, determining the flight altitude and Mach number of each design point corresponding to the flight mission;

[0014] Step S12: Calculate the corresponding engine inlet temperature and engine inlet pressure based on the flight altitude and Mach number.

[0015] Preferably, in step S2, the compressor outlet gas parameters include compressor outlet temperature, compressor outlet pressure, and compressor outlet flow rate, the main channel turbine rotor front gas parameters include high-pressure turbine rotor inlet temperature, high-pressure turbine rotor inlet pressure, and high-pressure turbine rotor inlet flow rate, and in step S3, the cooling gas parameters include cooling gas temperature, cooling gas pressure, and cooling gas flow rate.

[0016] Preferably, in step S3, the flow path parameters of the cooling system include the flow path size and flow path loss coefficient of the cooling system.

[0017] Preferably, in step S4, the theoretical turbine matrix temperature is calculated by a turbine three-dimensional simulation calculation program.

[0018] Preferably, in step S5, adjusting the initial heating temperature of the test includes:

[0019] When the ground turbine base temperature is greater than the theoretical turbine base temperature, lowering the initial heating temperature, and recalculating the ground turbine base temperature until the ground turbine base temperature is the same as the theoretical turbine base temperature;

[0020] When the ground turbine matrix temperature is lower than the theoretical turbine matrix temperature, the initial heating temperature is increased, and the ground turbine matrix temperature is recalculated until the ground turbine matrix temperature is the same as the theoretical turbine matrix temperature.

[0021] This application ensures that the damage during the endurance test fully simulates the actual damage during field use, which can fully evaluate the durability of the engine without causing excessive evaluation, thereby effectively ensuring flight safety. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 It is a flow chart of a preferred embodiment of the method for determining the ground intake heating temperature of an aircraft engine endurance test of the present application.

[0023] Figure 2 This is a schematic diagram of the engine gas path cross-section numbering. DETAILED DESCRIPTION

[0024] In order to make the purpose, technical scheme and advantages of the implementation of this application clearer, the technical scheme in the implementation of this application will be described in more detail in combination with the drawings in the implementation of this application. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and cannot be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in combination with the drawings.

[0025] The present application provides a method for determining the ground intake air heating temperature of an aircraft engine during a long-term test run, such as Figure 1 As shown, it mainly includes:

[0026] Step S1, determining the engine inlet temperature and the engine inlet pressure at each design point corresponding to the flight mission simulated by the aircraft engine endurance test;

[0027] Step S2, determining the compressor outlet gas parameters and the main flow channel turbine rotor front gas parameters according to the engine inlet temperature and the engine inlet pressure;

[0028] Step S3, calculating the cooling air parameters introduced into the turbine rotor according to the compressor outlet parameters and the flow path parameters of the cooling system for cooling the turbine rotor through the compressor bleed air;

[0029] Step S4, calculating the theoretical turbine base temperature according to the cooling gas parameters and the gas parameters before the main flow channel turbine rotor;

[0030] Step S5, determining the ground turbine base temperature calculated under the given ground normal pressure and initial heating temperature conditions during the aircraft engine endurance test;

[0031] Step S6: For each design point, the initial heating temperature of the test is adjusted according to the theoretical turbine matrix temperature and the ground turbine matrix temperature, so that the theoretical turbine matrix temperature is the same as the ground turbine matrix temperature, thereby forming the initial heating temperature of each design point of the aircraft engine endurance test.

[0032] In the present application, the main assessment in the endurance test is creep thermal damage, and the most direct parameter characterizing this damage is the turbine component substrate temperature. Under normal pressure conditions on the ground, by setting a suitable intake heating temperature to ensure that the turbine substrate temperature in the ground environment is equal to the turbine substrate temperature in actual flight, the damage mainly assessed in the endurance test can be made consistent with that in actual flight.

[0033] First, in step S1, this application mainly conducts endurance tests on aircraft engines, mainly simulating the durability of the hot end components of the engine under different flight mission conditions on the ground. Therefore, it is necessary to calculate the corresponding engine inlet temperature under each flight mission profile, that is, the initial heating temperature. First of all, it should be noted that the hot end components are key components that restrict the life of the engine, which can usually be determined through life damage assessment. This application takes the high-pressure turbine as an example. Secondly, this application determines multiple design points of the flight mission profile. For each design point, in order to calculate the initial heating temperature, the corresponding engine inlet parameters of the engine are first calculated, that is, the engine inlet temperature and the engine inlet pressure.

[0034] In some optional implementations, step S1 further includes:

[0035] Step S11, determining the flight altitude and Mach number of each design point corresponding to the flight mission;

[0036] Step S12: Calculate the corresponding engine inlet temperature and engine inlet pressure based on the flight altitude and Mach number.

[0037] Figure 2 The numbers of the gas path sections of the engine are given. For example, the number on the far left of the entire engine is 1, which means the engine inlet temperature is T1 and the engine inlet pressure is P1. The number before the compressor is 25, which means the compressor inlet temperature is T 25 , the compressor inlet pressure is P 25 .

[0038] Then in step S2, according to the engine control law, the overall engine performance parameters are calculated, including the compressor outlet gas parameters and the main flow channel turbine rotor front gas parameters.

[0039] In some optional embodiments, in step S2, the compressor outlet gas parameters include compressor outlet temperature T3, compressor outlet pressure P3, compressor outlet flow rate W3, and the main flow channel turbine rotor front gas parameters include high pressure turbine rotor inlet temperature T 41 , High-pressure turbine rotor inlet pressure P 41 and the high-pressure turbine rotor inlet flow rate W 41 .

[0040] Then in step S3, since the engine cools the turbine through the compressor bleed air, after the compressor outlet parameters are given, the cooling air parameters introduced into the turbine rotor can be calculated through the air path parameters of the cooling system.

[0041] In some optional embodiments, in step S3, the cooling gas parameters include the cooling gas temperature T 1q , Cooling gas pressure P 1q And cooling air flow W 1q .

[0042] In some optional embodiments, the flow path parameters of the cooling system include the flow path size and the flow path loss coefficient of the cooling system.

[0043] Then, in step S4, the theoretical temperature of the turbine base body under air conditions is calculated under the combined effect of the gas parameters before the main flow channel turbine rotor and the cooling gas parameters, that is, the theoretical turbine base body temperature T w1jt空中 .

[0044] In some optional embodiments, in step S4, the theoretical turbine matrix temperature is calculated by a turbine three-dimensional simulation calculation program.

[0045] At the same time, based on the calculation model of the above steps S2 to S4, in step S5, the temperature at the ground normal pressure and the initial heating temperature T is calculated. 1地面 Ground turbine base temperature T under conditions w1jt地面 .

[0046] Then in step S6, based on the ground turbine base temperature T w1jt地面 and the theoretical turbine base temperature T w1jt空中 Initial heating temperature T 1地面 Make adjustments.

[0047] In some optional embodiments, in step S5, adjusting the initial heating temperature of the test includes:

[0048] When the ground turbine base temperature is greater than the theoretical turbine base temperature, lowering the initial heating temperature, and recalculating the ground turbine base temperature until the ground turbine base temperature is the same as the theoretical turbine base temperature;

[0049] When the ground turbine matrix temperature is lower than the theoretical turbine matrix temperature, the initial heating temperature is increased, and the ground turbine matrix temperature is recalculated until the ground turbine matrix temperature is the same as the theoretical turbine matrix temperature.

[0050] Understandably, if T wljt地面 >T wljt空中 , it means that the heat load in ground heating is too heavy compared with the actual flight, and the initial heating temperature T needs to be lowered1地面 ; If T wljt地面 <T wljt空中 , it means that the heat load in ground heating is lower than that in actual flight, and the initial heating temperature T needs to be increased. 1地面 For each design point in the flight mission to be simulated, the initial heating temperature T of the corresponding engine air inlet is calculated according to the above method. 1地面 , the total intake air heating temperature in the endurance test can be obtained to form a temperature load spectrum.

[0051] This application ensures that the damage during the endurance test fully simulates the actual damage during field use, which can fully evaluate the durability of the engine without causing excessive evaluation, thereby effectively ensuring flight safety.

[0052] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A method for determining the ground intake air heating temperature of an aircraft engine endurance test, characterized in that: include: Step S1, determining the engine inlet temperature and the engine inlet pressure at each design point corresponding to the flight mission simulated by the aircraft engine endurance test; Step S2, determining the compressor outlet gas parameters and the main flow channel turbine rotor front gas parameters according to the engine inlet temperature and the engine inlet pressure; Step S3, calculating the cooling air parameters introduced into the turbine rotor according to the compressor outlet parameters and the flow path parameters of the cooling system for cooling the turbine rotor through the compressor bleed air; Step S4, calculating the theoretical turbine base temperature according to the cooling gas parameters and the gas parameters before the main flow channel turbine rotor; Step S5, determining the ground turbine base temperature calculated under the given ground normal pressure and initial heating temperature conditions during the aircraft engine endurance test; Step S6, for each design point, adjusting the initial heating temperature of the test according to the theoretical turbine base temperature and the ground turbine base temperature, so that the theoretical turbine base temperature is the same as the ground turbine base temperature, thereby forming the initial heating temperature of each design point of the aircraft engine endurance test; Wherein, step S1 further comprises: Step S11, determining the flight altitude and Mach number of each design point corresponding to the flight mission; Step S12: Calculate the corresponding engine inlet temperature and engine inlet pressure based on the flight altitude and Mach number.

2. The method for determining the ground intake air heating temperature of an aircraft engine endurance test according to claim 1, characterized in that: In step S2, the compressor outlet gas parameters include compressor outlet temperature, compressor outlet pressure, and compressor outlet flow rate, the main channel turbine rotor front gas parameters include high-pressure turbine rotor inlet temperature, high-pressure turbine rotor inlet pressure, and high-pressure turbine rotor inlet flow rate, and in step S3, the cooling gas parameters include cooling gas temperature, cooling gas pressure, and cooling gas flow rate.

3. The method for determining the ground intake air heating temperature of an aircraft engine endurance test as claimed in claim 1, characterized in that: In step S3, the flow path parameters of the cooling system include the flow path size and flow path loss coefficient of the cooling system.

4. The method for determining the ground intake air heating temperature of an aircraft engine endurance test as claimed in claim 1, characterized in that: In step S4, the theoretical turbine substrate temperature is calculated by a turbine three-dimensional simulation calculation program.

5. The method for determining the ground intake air heating temperature of an aircraft engine endurance test as claimed in claim 1, characterized in that: In step S5, adjusting the initial heating temperature of the test includes: When the ground turbine base temperature is greater than the theoretical turbine base temperature, lowering the initial heating temperature, and recalculating the ground turbine base temperature until the ground turbine base temperature is the same as the theoretical turbine base temperature; When the ground turbine matrix temperature is lower than the theoretical turbine matrix temperature, the initial heating temperature is increased, and the ground turbine matrix temperature is recalculated until the ground turbine matrix temperature is the same as the theoretical turbine matrix temperature.

Images