A method for upgrading an overall design of an aeroengine component

Abstract

This invention pertains to the field of engine technology and proposes an overall design method for upgrading aero-engine components. The method includes: ensuring the ratio of intake loss to reference intake pressure loss is within a preset range; meeting the performance requirements of the compressor (pressure performance), the combustion-related performance of the combustion chamber (combustion-related performance), the rotational inertia of the gas turbine (rotational inertia), the inlet temperature, and the performance requirements of the gas turbine; meeting the rotational inertia requirements of the reference power turbine and the turbine flow area requirements of the power turbine; and meeting the pressure loss requirements of the reference exhaust section and the exhaust velocity requirements of the exhaust section. This allows for the simultaneous or individual installation of improved components on a reference engine to enhance engine power and lifespan, among other technical indicators.

Description

Technical Field

[0001] This invention belongs to the field of engine technology, and specifically relates to an overall design method for upgrading aero-engine components. Background Technology

[0002] In the current serial development of aero engines, in order to improve the technical indicators such as power or lifespan of the development engine, the main components such as compressor, turbine, and tail nozzle are usually modified based on the reference engine. However, the compatibility between the modified components and the reference engine is usually not considered in the modification design, which means that the modified components cannot be installed and tested on the reference engine, and the reference engine cannot improve its performance by replacing a single modified component. Summary of the Invention

[0003] To address the above problems, this invention proposes an overall design method for upgrading aero-engine components, the method comprising:

[0004] The dimensions of multiple components each satisfy the condition that they are the same as the dimensions of multiple corresponding reference components;

[0005] One of the components is the intake section, and the ratio of the intake loss of the intake section to the pressure loss of the reference intake section is within a preset range;

[0006] The other component is a compressor, whose air pressure performance meets the performance requirements of a reference compressor;

[0007] The other component is the combustion chamber, whose combustion-related performance meets the performance conditions of a benchmark combustion chamber;

[0008] The other component is a gas turbine, whose moment of inertia meets the reference gas turbine moment of inertia condition, whose inlet temperature meets the reference gas turbine inlet temperature condition, and whose performance conditions are met.

[0009] Another component is a power turbine, whose rotational inertia satisfies the reference power turbine rotational inertia and the turbine flow area condition of the power turbine.

[0010] Another component is the exhaust section, whose pressure loss meets the pressure loss condition of the reference exhaust section, and whose exhaust velocity meets the exhaust speed condition.

[0011] Optionally, the ratio of the intake section intake loss to the reference intake section pressure loss is within a preset range, including:

[0012] The ratio of the intake loss in the intake section to the pressure loss in the reference intake section should be between 0.9 and 1.1.

[0013] Optionally, the compressor's pressure performance meets the performance conditions of a reference compressor, including:

[0014] The ratio of the compressor moment of inertia to the reference compressor moment of inertia should be controlled within a preset range.

[0015] Compressor outlet flow rate conversion:

[0016]

[0017] The compressor inlet flow rate is The boost ratio is Efficiency The reference compressor inlet flow rate is The reference boost ratio is Reference efficiency , is the specific heat ratio of the gas;

[0018] The surge margin of the compressor at each speed should not be lower than the surge margin of the reference compressor.

[0019] Optionally, the combustion-related performance of the combustion chamber meets the performance conditions of a benchmark combustion chamber, including:

[0020] The combustion efficiency of the combustion chamber is not lower than that of the reference combustion chamber;

[0021] The ratio of combustion chamber pressure loss to reference combustion chamber pressure loss is within a preset range;

[0022] The outlet temperature coefficient of the combustion chamber is not greater than that of the reference engine combustion chamber.

[0023] Optionally, the moment of inertia of the gas turbine satisfies the reference gas turbine moment of inertia condition, including:

[0024] When replacing the gas turbine on a reference engine, the ratio of the gas turbine's moment of inertia to that of the reference engine's gas turbine should be controlled within a preset range.

[0025] In addition, the compressor and gas turbine should be replaced simultaneously, and the ratio of the sum of the rotational inertia of the compressor and gas turbine to the sum of the rotational inertia of the compressor and gas turbine of the reference engine should be controlled within a preset range.

[0026] Optionally, the gas turbine inlet temperature meets the reference gas turbine inlet temperature conditions, including:

[0027] When replacing the gas turbine on a reference engine, the turbine inlet temperature should be the same as that of the reference engine turbine inlet temperature.

[0028] Optionally, the performance conditions for gas turbines include:

[0029] To replace both the compressor and the gas turbine simultaneously, the following conditions must be met:

[0030]

[0031] The compressor inlet flow rate is The boost ratio is Turbine inlet temperature is Reference engine compressor intake flow rate Boost ratio Turbine inlet temperature .

[0032] Optionally, the moment of inertia of the power turbine satisfies the reference moment of inertia of the power turbine, including:

[0033] The ratio of the moment of inertia of the power turbine to that of the reference engine power turbine should be controlled between 0.9 and 1.1.

[0034] Optionally, the turbine flow area conditions for the power turbine include:

[0035] Replace the power turbine on the base engine separately, keeping the flow area of ​​the power turbine the same as that of the base engine's power turbine;

[0036] At the same time, replace the compressor or gas turbine, and determine the new power turbine flow area according to the design procedure.

[0037] Optionally, the pressure loss of the exhaust section satisfies the pressure loss condition of the reference exhaust section, and the exhaust velocity condition of the exhaust section, including:

[0038] The pressure loss in the exhaust section is no greater than the pressure loss in the reference exhaust section.

[0039] The engine exhaust speed should not be less than the reference engine exhaust speed.

[0040] The overall design method for upgraded aero-engine components provided by this invention has the following advantages compared with existing technologies:

[0041] In addition to meeting the same conditions as multiple corresponding reference component dimensions, the design performance of each component must also be met. Specifically, the performance conditions of the intake section, compressor, combustion chamber, gas turbine, power turbine, and exhaust section must be met. This allows some components to be installed on the reference engine simultaneously to improve engine power and lifespan, while a single improved component can be installed on the reference engine to improve its power or lifespan.

[0042] Other features and advantages of the invention will be set forth in the following description, and will be apparent in part from the description, or may be learned by practicing the invention. The objects and other advantages of the invention can be realized and obtained by means of the structures pointed out in the description and the drawings. Attached Figure Description

[0043] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0044] Figure 1 A schematic flowchart of the overall design method for upgrading aero-engine components in an embodiment of the present invention is shown. Detailed Implementation

[0045] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0046] like Figure 1 As shown, the present invention provides an overall design method for upgrading aero-engine components, the method comprising:

[0047] Step S10: The dimensions of multiple parts respectively meet the condition that they are the same as the dimensions of multiple corresponding reference parts.

[0048] It should be noted that the axial length, inlet size, and outlet size of the intake section should be consistent with the reference intake section, and the maximum profile should not exceed that of the reference intake section. The axial length, inlet size, outlet size, and helicopter bleed air position of the compressor should be consistent with the reference compressor, and the maximum profile should not exceed that of the reference compressor. The axial length, inlet size, and outlet size of the combustion chamber should be consistent with the reference engine combustion chamber, and the maximum profile should not exceed that of the reference engine combustion chamber. The axial length, inlet size, and outlet size of the gas turbine should be consistent with the reference engine gas turbine, and the maximum profile should not exceed that of the reference engine gas turbine. The axial length, inlet size, and outlet size of the power turbine should be consistent with the reference engine power turbine, and the maximum profile should not exceed that of the reference engine power turbine. The inlet size of the exhaust section should be consistent with the reference exhaust section, and the overall length and outlet size can be determined according to the interface size of the helicopter platform.

[0049] In step S20, one of the components is the intake section, and the ratio of the intake loss of the intake section to the pressure loss of the reference intake section is within a preset range. This allows for the replacement of the intake section.

[0050] In one embodiment, the ratio of intake section intake loss to reference intake section pressure loss is within a preset range, including:

[0051] The ratio of intake loss in the intake section to pressure loss in the reference intake section should be between 0.9 and 1.1. The ratio of pressure loss should also be between 0.9 and 1.1 to ensure the matching characteristics of the engine air system.

[0052] In step S30, the other component is the compressor, whose pressure performance meets the performance requirements of the reference compressor. This allows the compressor to be replaced.

[0053] In one embodiment, the compressor's pressure performance meets the performance conditions of a reference compressor, including:

[0054] The ratio of the compressor's moment of inertia to the reference compressor's moment of inertia should be controlled within a preset range. Specifically, the ratio should be controlled between 0.95 and 1.05. If the ratio exceeds this range, the fuel control plan for engine starting and acceleration / deceleration needs to be adjusted accordingly. This adjustment can be achieved by controlling the fuel output.

[0055] Compressor outlet flow rate conversion:

[0056]

[0057] The compressor inlet flow rate is The boost ratio is Efficiency The reference compressor inlet flow rate is The reference boost ratio is Reference efficiency , Let be the specific heat ratio of the gas. The above flow rate formula is used to ensure that the converted flow rate at the compressor outlet remains constant.

[0058] The surge margin of the compressor at all speeds should not be lower than the surge margin of the reference engine compressor.

[0059] In step S40, the other component is the combustion chamber, whose combustion-related performance meets the performance conditions of the reference combustion chamber. This allows the combustion chamber to be replaced.

[0060] In one embodiment, the combustion-related performance of the combustion chamber meets the performance conditions of a reference combustion chamber, including:

[0061] The combustion efficiency of the combustion chamber is not lower than that of the reference combustion chamber to ensure the engine power characteristics.

[0062] The ratio of combustion chamber pressure loss to reference combustion chamber pressure loss is within a preset range. Specifically, the pressure loss ratio is between 0.9 and 1.1 to ensure the matching characteristics of the engine air system.

[0063] The outlet temperature coefficient of the combustion chamber is not greater than that of the reference engine combustion chamber to ensure the service life of the gas turbine blades.

[0064] In step S50, the other component is a gas turbine. The rotational inertia of the gas turbine meets the reference gas turbine rotational inertia condition, the gas turbine inlet temperature meets the reference gas turbine inlet temperature condition, and the performance conditions of the gas turbine are also met. This achieves the successful upgrade and replacement of the gas turbine.

[0065] In one embodiment, the moment of inertia of the gas turbine satisfies the reference gas turbine moment of inertia condition, including:

[0066] When replacing the gas turbine on a reference engine, the ratio of the gas turbine's moment of inertia to that of the reference engine's gas turbine should be controlled within a preset range. This ensures that the engine's start-up and acceleration / deceleration times meet the requirements. It should be noted that the ratio of the gas turbine's moment of inertia to that of the reference engine's gas turbine should be controlled between 0.95 and 1.05.

[0067] Simultaneously replace the compressor and gas turbine. The ratio of the sum of the compressor and gas turbine's moments of inertia to the sum of the moments of inertia of the reference engine's compressor and gas turbine should be controlled within a preset range. Optionally, the ratio should be controlled between 0.95 and 1.05. If the ratio exceeds the above range, the fuel control plan during engine start-up and acceleration / deceleration needs to be adjusted accordingly, for example, by increasing or decreasing combustion output.

[0068] In one embodiment, the gas turbine inlet temperature satisfies a reference gas turbine inlet temperature condition, including:

[0069] When replacing the gas turbine on a reference engine, the turbine inlet temperature should be the same as that of the reference engine. This allows for the successful replacement of the gas turbine independently.

[0070] In one embodiment, satisfying the performance conditions of the gas turbine includes:

[0071] To replace both the compressor and the gas turbine simultaneously, the following conditions must be met:

[0072]

[0073] The compressor inlet flow rate is The boost ratio is Turbine inlet temperature is Reference engine compressor intake flow rate Boost ratio Turbine inlet temperature By satisfying the above formula conditions, it is possible to replace both the compressor and the gas turbine simultaneously.

[0074] In step S60, the other component is the power turbine. The rotational inertia of the power turbine satisfies the rotational inertia of the reference power turbine and also meets the turbine flow area requirements of the power turbine. This allows the power turbine to be replaced.

[0075] In one embodiment, the moment of inertia of the power turbine satisfies the reference moment of inertia of the power turbine, including:

[0076] The ratio of the moment of inertia of the power turbine to that of the reference engine's power turbine should be controlled between 0.9 and 1.1. This ensures the engine's operational stability under no-load conditions.

[0077] In one embodiment, satisfying the turbine flow area condition of the power turbine includes:

[0078] The power turbine is replaced separately on the reference engine, ensuring that the flow area of ​​the power turbine is the same as that of the reference engine's power turbine. This guarantees the success of the separate power turbine replacement.

[0079] Simultaneously, replace the compressor or gas turbine, and determine the new power turbine flow area according to the design procedure. It should be noted that the determined new power turbine flow area must meet the performance requirements for replacement on the benchmark engine.

[0080] In step S70, another component is the exhaust section. The pressure loss of the exhaust section meets the pressure loss condition of the reference exhaust section, as well as the exhaust velocity condition of the exhaust section. This allows the exhaust section to be replaced.

[0081] In one embodiment, the pressure loss of the exhaust section satisfies the pressure loss condition of the reference exhaust section and the exhaust velocity condition of the exhaust section, including:

[0082] The pressure loss in the exhaust section is no greater than the pressure loss in the reference exhaust section.

[0083] The engine exhaust velocity should not be less than the reference engine exhaust velocity. Specifically, the engine exhaust velocity should not be less than the reference engine exhaust velocity:

[0084]

[0085] The compressor inlet flow rate is The exhaust section outlet area is The exhaust temperature is The reference engine compressor intake flow rate is The exhaust section outlet area is The exhaust temperature is .

[0086] By improving the design of components such as compressors, turbines, and intake and exhaust systems, these components can be simultaneously installed on a reference engine to improve its power and lifespan, or a single improved component can be installed on the reference engine to improve its power or lifespan. The ability to install individual improved components, either in groups or individually, on reference engines to enhance their performance or lifespan significantly increases their versatility and economic benefits.

[0087] Although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for the overall design of upgraded aero-engine components, characterized in that, The method includes: The dimensions of multiple components each satisfy the condition that they are the same as the dimensions of multiple corresponding reference components; One of the components is the intake section, and the ratio of the intake loss of the intake section to the pressure loss of the reference intake section is within a preset range; The other component is a compressor, whose air pressure performance meets the performance requirements of a reference compressor; The other component is the combustion chamber, whose combustion-related performance meets the performance conditions of a benchmark combustion chamber; The other component is a gas turbine, whose moment of inertia meets the reference gas turbine moment of inertia condition, whose inlet temperature meets the reference gas turbine inlet temperature condition, and whose performance conditions are met. Another component is a power turbine, whose rotational inertia satisfies the reference power turbine rotational inertia and the turbine flow area condition of the power turbine. The other component is the exhaust section, whose pressure loss meets the pressure loss condition of the reference exhaust section and the exhaust velocity corresponding condition of the exhaust section; The ratio of the intake section intake loss to the reference intake section pressure loss is within a preset range, including: The ratio of the intake section intake loss to the reference intake section pressure loss should be between 0.9 and 1.

1. The compressor's pressure performance meets the performance requirements of a reference compressor, including: The ratio of the moment of inertia of the compressor to that of the reference compressor should be controlled within a preset range. Compressor outlet flow rate conversion: The compressor inlet flow rate is The boost ratio is Efficiency The reference compressor inlet flow rate is The reference boost ratio is Reference efficiency , is the specific heat ratio of the gas; The surge margin of the compressor at each speed should not be lower than the surge margin of the reference compressor.

2. The overall design method for upgraded aero-engine components according to claim 1, characterized in that, The combustion-related performance of the combustion chamber meets the performance conditions of a benchmark combustion chamber, including: The combustion efficiency of the combustion chamber is not lower than that of the reference combustion chamber; The ratio of combustion chamber pressure loss to reference combustion chamber pressure loss is within a preset range; The outlet temperature coefficient of the combustion chamber is not greater than that of the reference engine combustion chamber.

3. The overall design method for upgraded aero-engine components according to claim 1, characterized in that, The moment of inertia of the gas turbine satisfies the reference gas turbine moment of inertia condition, including: When replacing the gas turbine on a reference engine, the ratio of the gas turbine's moment of inertia to that of the reference engine's gas turbine should be controlled within a preset range. In addition, the compressor and gas turbine should be replaced simultaneously, and the ratio of the sum of the rotational inertia of the compressor and gas turbine to the sum of the rotational inertia of the compressor and gas turbine of the reference engine should be controlled within a preset range.

4. The overall design method for upgraded aero-engine components according to claim 1, characterized in that, The gas turbine inlet temperature meets the reference gas turbine inlet temperature conditions, including: When replacing the gas turbine on a reference engine, the turbine inlet temperature should be the same as that of the reference engine turbine inlet temperature.

5. The overall design method for upgraded aero-engine components according to claim 1, characterized in that, The performance conditions for gas turbines include: To replace both the compressor and the gas turbine simultaneously, the following conditions must be met: The compressor inlet flow rate is The boost ratio is Turbine inlet temperature is Reference engine compressor intake flow rate Boost ratio Turbine inlet temperature .

6. The overall design method for upgrading aero-engine components according to claim 1, characterized in that, The moment of inertia of the power turbine satisfies the reference moment of inertia of the power turbine, including: The ratio of the moment of inertia of the power turbine to that of the reference engine power turbine should be controlled between 0.9 and 1.

1.

7. The overall design method for upgrading aero-engine components according to claim 1, characterized in that, The turbine flow area conditions for the power turbine include: Replace the power turbine on the base engine separately, keeping the flow area of ​​the power turbine the same as that of the base engine's power turbine; At the same time, replace the compressor or gas turbine, and determine the new power turbine flow area according to the design procedure.

8. The overall design method for upgraded aero-engine components according to claim 1, characterized in that, The pressure loss of the exhaust section satisfies the pressure loss condition of the reference exhaust section, and also satisfies the exhaust velocity corresponding condition of the exhaust section, including: The pressure loss in the exhaust section is no greater than the pressure loss in the reference exhaust section. The engine exhaust speed should not be less than the reference engine exhaust speed.

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