A three spool multi-stage combustion turbine engine and method of operating the same

Abstract

The application belongs to the technical field of turbine engine design, and particularly relates to a three-rotor multi-stage combustion turbine engine and a control method thereof. A regulating valve is designed to control the turbine engine to be in a small-bypass-ratio three-stage combustion mode when cruising or taking off at a low Mach number, and to be in a low-pressure-ratio two-stage combustion mode when accelerating or cruising at a high Mach number, so that the thrust and efficiency of the turbine engine when cruising or taking off at a low Mach number and when accelerating or cruising at a high Mach number can be effectively ensured, the difficulty of engine thermal management can be reduced, and the connection with ram pressure can be ensured.

Description

Technical Field

[0001] This application belongs to the field of turbine engine design technology, specifically relating to a three-rotor multi-stage combustion turbine engine and its control method. Background Technology

[0002] Aerospace propulsion typically employs a combination of turbine engines and ramjet engines to meet the needs of different speed ranges. Turbine engines have higher efficiency at low Mach numbers and mainly provide power at low Mach numbers, while ramjet engines have higher efficiency at high Mach numbers and mainly provide power at high Mach numbers.

[0003] Existing twin-rotor turbine engines expose the core to high-temperature incoming airflow during high Mach number acceleration and cruise, posing significant challenges to engine thermal management and the application of high-temperature resistant materials. Furthermore, the airflow mainly enters the afterburner through the outer bypass duct, resulting in low combustion efficiency and a risk of insufficient connection with ramjet thrust.

[0004] Existing three-rotor turbine engines with interstage combustion expose the core engine to high-temperature incoming flow during high Mach number acceleration and cruise, posing significant challenges to engine thermal management and the application of high-temperature resistant materials. Furthermore, the lack of necessary adjustment mechanisms makes it difficult to fully utilize the synergistic effect of the interstage combustion chamber and afterburner, and also carries the risk of insufficient ramjet thrust.

[0005] This application is made in view of the aforementioned technical deficiencies. Summary of the Invention

[0006] The purpose of this application is to provide a three-rotor multi-stage combustion turbine engine and its control method to overcome or mitigate at least one of the known technical defects.

[0007] The technical solution of this application is:

[0008] One aspect provides a three-rotor multi-stage combustion turbine engine, comprising:

[0009] Low-pressure fan;

[0010] The medium-pressure compressor is installed after the low-pressure fan;

[0011] The high-pressure compressor is installed after the medium-pressure compressor.

[0012] The main combustion chamber is located after the high-pressure compressor;

[0013] The high-pressure turbine is located after the main combustion chamber and is connected to the high-pressure compressor via a high-pressure shaft.

[0014] The intermediate-pressure interstage combustion chamber is located after the high-pressure turbine and is connected to the intermediate-pressure compressor via a first external bypass duct;

[0015] The intermediate-pressure turbine is located after the intermediate-pressure stage combustion chamber and is connected to the intermediate-pressure compressor via an intermediate-pressure shaft.

[0016] The low-pressure turbine is installed after the medium-pressure turbine and is connected to the low-pressure fan via a low-pressure shaft.

[0017] An afterburner is located after the low-pressure turbine, connected to the intermediate-pressure compressor via a first bypass duct and to the low-pressure fan via a second bypass duct; the first and second bypass ducts are separated by a split ring.

[0018] The nozzle is located after the afterburner.

[0019] The second outer bypass valve is connected after the low-pressure fan. When closed, it is biased towards the inner flow channel surface of the second outer bypass, cutting off the flow path between the afterburner and the low-pressure fan; when opened, it is biased towards the outer flow channel surface of the second outer bypass, allowing the flow path between the afterburner and the low-pressure fan to be connected.

[0020] The valve before the first outer duct is connected to the high-pressure compressor. When closed, it is biased towards the high-pressure compressor, reducing the flow area between the high-pressure compressor and the intermediate-pressure compressor; when open, it is biased towards the flow channel surface inside the first outer duct, increasing the flow area between the high-pressure compressor and the intermediate-pressure compressor.

[0021] The valve after the first bypass duct is connected after the high-pressure turbine. When closed, it is biased towards the flow channel surface inside the first bypass duct, increasing the flow area between the high-pressure turbine and the intermediate-pressure stage combustion chamber, and shutting off the flow path between the intermediate-pressure stage combustion chamber and the intermediate-pressure compressor. When open, it is biased towards the high-pressure turbine, reducing the flow area between the high-pressure turbine and the intermediate-pressure stage combustion chamber, and connecting the flow path between the intermediate-pressure stage combustion chamber and the intermediate-pressure compressor.

[0022] The downstream bypass valve is connected after the flow divider ring. When closed, it deflects towards the inner flow channel of the first outer bypass, shutting off the flow path between the afterburner and the intermediate-pressure compressor; when open, it deflects towards the outer flow channel of the second outer bypass, shutting off the flow path between the afterburner and the low-pressure fan.

[0023] According to at least one embodiment of this application, in the above-described three-rotor multi-stage combustion turbine engine, the three-rotor multi-stage combustion turbine engine has:

[0024] In the low bypass ratio three-stage combustion mode, the second bypass valve is closed, the first bypass front valve is open, the first bypass rear valve is closed, the rear bypass valve is open, and the main combustion chamber is working.

[0025] In the low-pressure two-stage combustion mode, the second outer duct valve is open, the first outer duct front valve is closed, the first outer duct rear valve is open, the rear duct valve is closed, and the main combustion chamber does not work.

[0026] According to at least one embodiment of this application, in the above-described three-rotor multi-stage combustion turbine engine, the high-pressure shaft is sleeved on the outer periphery of the intermediate-pressure shaft and supported therebetween by bearings.

[0027] The intermediate pressure shaft is sleeved on the outer circumference of the low pressure shaft, and is supported by bearings in between.

[0028] According to at least one embodiment of this application, in the above-described three-rotor multi-stage combustion turbine engine, the intermediate-pressure turbine is an adjustable turbine.

[0029] According to at least one embodiment of this application, in the above-described three-rotor multi-stage combustion turbine engine, the low-pressure turbine is an adjustable turbine.

[0030] According to at least one embodiment of this application, in the above-described three-rotor multi-stage combustion turbine engine, the nozzle is an adjustable nozzle.

[0031] According to at least one embodiment of this application, the above-described three-rotor multi-stage combustion turbine engine further includes:

[0032] The mixing chamber is located at the end of the afterburner.

[0033] On the other hand, a method for controlling a three-rotor multi-stage combustion turbine engine is provided, for controlling the aforementioned three-rotor multi-stage combustion turbine engine, including:

[0034] When cruising at low Mach numbers or taking off, the three-rotor multi-stage combustion turbine engine is set to a low bypass ratio three-stage combustion mode, the second bypass valve is closed, the first bypass front valve is open, the first bypass rear valve is closed, the rear bypass valve is open, and the main combustion chamber is working.

[0035] When accelerating or cruising at high Mach numbers, the three-rotor multi-stage combustion turbine engine is set to a low-pressure ratio two-stage combustion mode, the second bypass valve is open, the first bypass front valve is closed, the first bypass rear valve is open, the rear bypass valve is closed, and the main combustion chamber is not working.

[0036] This application has at least the following beneficial technical effects:

[0037] This invention provides a three-rotor multi-stage combustion turbine engine and its control method. The design includes a regulating valve that controls the turbine engine to operate in a low-bypass ratio three-stage combustion mode during low Mach number cruise or takeoff, and in a low-pressure ratio two-stage combustion mode during high Mach number acceleration or cruise. This effectively ensures the thrust and efficiency of the turbine engine during low Mach number cruise or takeoff, as well as high Mach number acceleration or cruise, while also reducing the difficulty of engine thermal management and ensuring seamless integration with ramjet thrust. Attached Figure Description

[0038] Figure 1This is a schematic diagram of a three-rotor multi-stage combustion turbine engine mechanism provided in an embodiment of this application;

[0039] Figure 2 This is a schematic diagram of a three-rotor multi-stage combustion turbine engine provided in an embodiment of this application in a low-bypass ratio three-stage combustion mode;

[0040] Figure 3 This is a schematic diagram of the low bypass ratio three-stage combustion mode in the low pressure ratio two-stage combustion mode provided in the embodiments of this application;

[0041] in:

[0042] 1-Low-pressure fan; 2-Intermediate-pressure compressor; 3-High-pressure compressor; 4-Main combustion chamber; 5-High-pressure turbine; 6-Intermediate-pressure interstage combustion chamber; 7-Intermediate-pressure turbine; 8-Low-pressure turbine; 9-Afterburner; 10-Nozzle; 11-Second duct valve; 12-First duct inlet valve; 13-First duct outlet valve; 14-Rear duct valve; 15-Mixing chamber.

[0043] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual product dimensions. Furthermore, the drawings are for illustrative purposes only and should not be construed as limiting this application. Detailed Implementation

[0044] 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.

[0045] 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 terms indicating direction used in this application description are used only to indicate relative direction or positional relationship; when the absolute position of the described object changes, its relative positional relationship may also change accordingly. The word "comprising" as used in this application description indicates that the element or object preceding the word encompasses the elements or objects listed following the word and their equivalents, but does not exclude other elements or objects.

[0046] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0047] On the one hand, a three-rotor multi-stage combustion turbine engine is provided, such as Figure 1 As shown, it includes:

[0048] Low-pressure fan 1;

[0049] Medium-pressure compressor 2 is installed after low-pressure fan 1;

[0050] High-pressure compressor 3 is installed after medium-pressure compressor 2;

[0051] The main combustion chamber 4 is located after the high-pressure compressor 3;

[0052] The high-pressure turbine 5 is located after the main combustion chamber 4 and is connected to the high-pressure compressor 3 via a high-pressure shaft.

[0053] Intermediate-stage combustion chamber 6 is located after high-pressure turbine 5 and is connected to intermediate-pressure compressor 2 via first external bypass duct;

[0054] The intermediate-pressure turbine 7 is installed after the intermediate-pressure interstage combustion chamber 6 and is connected to the intermediate-pressure compressor 2 via an intermediate-pressure shaft.

[0055] The low-pressure turbine 8 is installed after the medium-pressure turbine 7 and is connected to the low-pressure fan 1 via a low-pressure shaft.

[0056] An afterburner 9 is located after the low-pressure turbine 8, and is connected to the intermediate-pressure compressor 2 through a first bypass duct and to the low-pressure fan 1 through a second bypass duct; the first bypass duct and the second bypass duct are separated by a split ring.

[0057] The nozzle 10 is located after the afterburner 9;

[0058] The second outer duct valve 11 is connected after the low-pressure fan 1. When closed, it is biased towards the inner flow channel surface of the second outer duct, cutting off the flow path between the afterburner 9 and the low-pressure fan 1; when opened, it is biased towards the outer flow channel surface of the second outer duct, connecting the flow path between the afterburner 9 and the low-pressure fan 1.

[0059] The first outer duct front valve 12 is connected to the front of the high-pressure compressor 3. When closed, it is biased towards the high-pressure compressor 3, reducing the flow area between the high-pressure compressor 3 and the intermediate-pressure compressor 2; when open, it is biased towards the flow channel surface inside the first outer duct, increasing the flow area between the high-pressure compressor 3 and the intermediate-pressure compressor 2.

[0060] The first outer bypass valve 13 is connected after the high-pressure turbine 5. When closed, it is biased towards the flow channel surface inside the first outer bypass duct, increasing the flow area between the high-pressure turbine 5 and the intermediate-pressure stage combustion chamber 6, and shutting off the flow path between the intermediate-pressure stage combustion chamber 6 and the intermediate-pressure compressor 2. When open, it is biased towards the high-pressure turbine 5, reducing the flow area between the high-pressure turbine 5 and the intermediate-pressure stage combustion chamber 6, and connecting the flow path between the intermediate-pressure stage combustion chamber 6 and the intermediate-pressure compressor 2.

[0061] The downstream bypass valve 14 is connected after the flow divider ring. When closed, it is biased towards the inner flow channel surface of the first outer bypass, shutting off the flow path between the afterburner 9 and the intermediate-pressure compressor 2; when open, it is biased towards the outer flow channel surface of the second outer bypass, shutting off the flow path between the afterburner 9 and the low-pressure fan 1.

[0062] The three-rotor multi-stage combustion turbine engine disclosed in the above embodiments may have:

[0063] In the low bypass ratio three-stage combustion mode, the second bypass valve 11 is closed, the first bypass front valve 12 is open, the first bypass rear valve 13 is closed, the rear bypass valve 14 is open, and the main combustion chamber is working.

[0064] In the low-pressure two-stage combustion mode, the second outer duct valve 11 is open, the first outer duct front valve 12 is closed, the first outer duct rear valve 13 is open, the rear duct valve 14 is closed, and the main combustion chamber does not work.

[0065] During low Mach number cruise or takeoff, the three-rotor multi-stage combustion turbine engine can be set to a low bypass ratio three-stage combustion mode. In this mode, high-temperature air enters the turbine engine, is compressed by the low-pressure fan 1, and then flows to the intermediate-pressure compressor 2. The airflow in the intermediate-pressure compressor 2 is further compressed and splits into two streams, flowing to the first outer bypass duct and the high-pressure compressor 3 respectively. The airflow to the high-pressure compressor 3 is further compressed and enters the main combustion chamber 4, where it mixes with fuel and burns to produce high-temperature gas. This high-temperature gas flows into the high-pressure turbine 5, expands, and performs work. Through the high-pressure shaft, it drives the high-pressure compressor 3. The high-temperature gas after expansion and work... The temperature decreases, and then it enters the intermediate-pressure interstage combustion chamber 6, where it is mixed with fuel again to produce high-temperature gas. This high-temperature gas enters the intermediate-pressure turbine 7 to expand and do work, driving the intermediate-pressure fan 2 through the intermediate-pressure shaft. After expanding and doing work, the temperature of the high-temperature gas decreases, and then it enters the low-pressure turbine 8 to expand and do work, driving the intermediate-pressure compressor 2 through the low-pressure shaft. After expanding and doing work, the temperature of the high-temperature gas decreases further and flows to the afterburner 9. In the afterburner 9, it mixes with high-temperature air from the first bypass duct and is mixed with fuel to produce high-temperature gas. After expanding through the nozzle 10, it is ejected to generate thrust. Figure 2As shown, in the low bypass ratio three-stage combustion mode, the airflow mainly burns through the inner main combustion chamber 4, the intermediate stage combustion chamber 6, and the inner and outer bypass mixture enters the afterburner 9 for afterburning. The three-stage combustion can significantly increase the propulsion power of the turbine engine, improve the thrust of the turbine engine at low Mach numbers, meet the thrust requirements for takeoff acceleration of large-tonnage hypersonic aircraft, and ensure the thrust for low Mach number cruise. Moreover, the interstage combustion has high efficiency, and the use of the first outer bypass air to supplement oxygen to the afterburner 9 can improve the afterburning effect. The first outer bypass air entering the afterburner 9 is compressed in two stages by the low-pressure fan 1 and the intermediate-pressure compressor 2, and has a high temperature and pressure, which can fully realize the potential of afterburning.

[0066] During high Mach number acceleration or cruising, the three-rotor multi-stage combustion turbine engine is set to a low-pressure ratio two-stage combustion mode. In this mode, high-temperature air enters the turbine engine, is compressed by the low-pressure fan 1, and is split into two streams, flowing to the second bypass duct and the intermediate-pressure compressor 2, respectively. The airflow to the intermediate-pressure compressor 2 is further compressed and split into two streams, flowing to the first bypass duct and the high-pressure compressor 3, respectively. The airflow to the high-pressure compressor 3 is further compressed and enters the main combustion chamber 4, then flows into the high-pressure turbine 5 for expansion and power generation. This expansion and power generation drives the high-pressure compressor 3 via the high-pressure shaft. The high-temperature air, after expansion and power generation, cools down and enters the intermediate-pressure interstage combustion chamber. Inside combustion chamber 6, the high-temperature air from the first bypass duct mixes with fuel and combusts to produce high-temperature gas. This high-temperature gas enters the intermediate-pressure turbine 7, expands, and performs work. It drives the intermediate-pressure fan 2 via the intermediate-pressure shaft. After expansion and work, the temperature of the high-temperature gas decreases. It then enters the low-pressure turbine 8, expands, and performs work again. It drives the intermediate-pressure compressor 2 via the low-pressure shaft. The temperature of the high-temperature gas further decreases after expansion and work, and it flows to the afterburner 9. Inside the afterburner 9, it mixes with high-temperature air from the second bypass duct and combusts to produce high-temperature gas. This gas expands through nozzle 10 and is then ejected, generating thrust. Figure 3 As shown, at this time, the main combustion chamber 4 is closed, and the core engine is less exposed to the high-temperature incoming flow, which reduces the difficulty of thermal management of the turbine engine and protects the turbine engine from high-temperature damage. Under the condition of high-temperature incoming flow, the air entering the intermediate-stage combustion chamber 6 is compressed in three stages by the low-pressure fan 1, intermediate-pressure compressor 2, and high-pressure compressor 3, and has a high temperature, which can ensure the efficiency of the intermediate-stage combustion chamber 6. In addition, the second bypass air with high pressure and temperature is used to supplement oxygen for the afterburner 9, which can improve the afterburning effect. The cooperation effect of the interstage combustion chamber 6 and the afterburner 9 can be fully utilized to improve the thrust of the turbine engine at high Mach numbers, realize the use of the turbine engine in a wide speed range, and ensure the connection of the thrust of the turbine-ramjet combined power system for high Mach number state.

[0067] In some alternative embodiments, in the above-described three-rotor multi-stage combustion turbine engine, the high-pressure shaft is sleeved on the outer periphery of the intermediate-pressure shaft and supported therebetween by bearings, and can be supported on the corresponding stator components by bearings.

[0068] The intermediate pressure shaft is sleeved on the outer circumference of the low pressure shaft, supported by bearings in between, and can also be supported by bearings on the corresponding stator components.

[0069] In some alternative embodiments, the intermediate-pressure turbine 7 in the above-described three-rotor multi-stage combustion turbine engine is an adjustable turbine.

[0070] In some alternative embodiments, the low-pressure turbine 8 in the above-described three-rotor multi-stage combustion turbine engine is an adjustable turbine.

[0071] In some alternative embodiments, the nozzle 10 in the above-described three-rotor multi-stage combustion turbine engine is an adjustable nozzle.

[0072] In some optional embodiments, the above-described three-rotor multi-stage combustion turbine engine further includes:

[0073] The mixing chamber 15 is located at the end of the afterburner 9.

[0074] The three-rotor multi-stage combustion turbine engine disclosed in the above embodiments can adopt a modular design concept, which facilitates low-cost and rapid maintenance and support as well as serial development.

[0075] On the other hand, a method for controlling a three-rotor multi-stage combustion turbine engine is provided, for controlling the aforementioned three-rotor multi-stage combustion turbine engine, including:

[0076] When cruising at low Mach numbers or taking off, the three-rotor multi-stage combustion turbine engine is set to a low bypass ratio three-stage combustion mode, the second bypass valve 11 is closed, the first bypass front valve 12 is open, the first bypass rear valve 13 is closed, the rear bypass valve 14 is open, and the main combustion chamber is working.

[0077] When accelerating or cruising at high Mach numbers, the three-rotor multi-stage combustion turbine engine is set to a low-pressure ratio two-stage combustion mode, the second bypass valve 11 is open, the first bypass front valve 12 is closed, the first bypass rear valve 13 is open, the rear bypass valve 14 is closed, and the main combustion chamber is not working.

[0078] The control method for the three-rotor multi-stage combustion turbine engine disclosed in the above embodiments is used to control the three-rotor multi-stage combustion turbine engine. The description is relatively simple. For specific details, please refer to the relevant description of the three-rotor multi-stage combustion turbine engine. The technical effects can also be referred to the technical effects of the relevant parts of the three-rotor multi-stage combustion turbine engine. It will not be repeated here.

[0079] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. Similar or identical parts between embodiments can be referred to mutually. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined to obtain new embodiments.

[0080] 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 three-rotor multi-stage combustion turbine engine, characterized in that, include: Low-pressure fan (1); A medium-pressure compressor (2) is installed after a low-pressure fan (1); A high-pressure compressor (3) is installed after the medium-pressure compressor (2); The main combustion chamber (4) is located after the high-pressure compressor (3); A high-pressure turbine (5) is installed after the main combustion chamber (4) and is connected to a high-pressure compressor (3) via a high-pressure shaft; Intermediate-stage combustion chamber (6) is located after high-pressure turbine (5) and is connected to intermediate-pressure compressor (2) via first external bypass duct; The intermediate-pressure turbine (7) is installed after the intermediate-pressure interstage combustion chamber (6) and is connected to the intermediate-pressure compressor (2) via the intermediate-pressure shaft; A low-pressure turbine (8) is installed after the medium-pressure turbine (7) and is connected to a low-pressure fan (1) via a low-pressure shaft; An afterburner (9) is located after the low-pressure turbine (8), and is connected to the intermediate-pressure compressor (2) through a first bypass duct and to the low-pressure fan (1) through a second bypass duct; the first bypass duct and the second bypass duct are separated by a split ring; The nozzle (10) is located after the afterburner (9); The second outer duct valve (11) is connected after the low-pressure fan (1). When closed, it is biased towards the inner flow channel surface of the second outer duct, cutting off the flow path between the afterburner (9) and the low-pressure fan (1); when opened, it is biased towards the outer flow channel surface of the second outer duct, connecting the flow path between the afterburner (9) and the low-pressure fan (1). The first outer duct front valve (12) is connected in front of the high pressure compressor (3). When closed, it is biased towards the high pressure compressor (3) to reduce the flow area between the high pressure compressor (3) and the medium pressure compressor (2); when opened, it is biased towards the flow channel surface inside the first outer duct to increase the flow area between the high pressure compressor (3) and the medium pressure compressor (2). The first outer bypass valve (13) is connected after the high-pressure turbine (5). When closed, it is biased towards the flow channel surface inside the first outer bypass duct, increasing the flow area between the high-pressure turbine (5) and the intermediate-pressure stage combustion chamber (6), and shutting off the flow path between the intermediate-pressure stage combustion chamber (6) and the intermediate-pressure compressor (2). When opened, it is biased towards the high-pressure turbine (5), reducing the flow area between the high-pressure turbine (5) and the intermediate-pressure stage combustion chamber (6), and connecting the flow path between the intermediate-pressure stage combustion chamber (6) and the intermediate-pressure compressor (2). The downstream bypass valve (14) is connected after the flow divider ring. When closed, it is biased towards the inner flow channel of the first outer bypass, shutting off the flow path between the afterburner (9) and the intermediate-pressure compressor (2); when open, it is biased towards the outer flow channel of the second outer bypass, shutting off the flow path between the afterburner (9) and the low-pressure fan (1).

2. The three-rotor multi-stage combustion turbine engine according to claim 1, characterized in that, The three-rotor multi-stage combustion turbine engine has the following characteristics: In the low bypass ratio three-stage combustion mode, the second bypass valve (11) is closed, the first bypass front valve (12) is open, the first bypass rear valve (13) is closed, the rear bypass valve (14) is open, and the main combustion chamber is working. In the low-pressure two-stage combustion mode, the second duct valve (11) is open, the first duct front valve (12) is closed, the first duct rear valve (13) is open, the rear duct valve (14) is closed, and the main combustion chamber does not work.

3. The three-rotor multi-stage combustion turbine engine according to claim 2, characterized in that, The high-pressure bushing is located on the outer circumference of the intermediate-pressure shaft, and is supported by bearings in between; The intermediate pressure shaft is sleeved on the outer circumference of the low pressure shaft, and is supported by bearings in between.

4. The three-rotor multi-stage combustion turbine engine according to claim 3, characterized in that, The medium-pressure turbine (7) is an adjustable turbine.

5. The three-rotor multi-stage combustion turbine engine according to claim 4, characterized in that, The low-pressure turbine (8) is an adjustable turbine.

6. The three-rotor multi-stage combustion turbine engine according to claim 5, characterized in that, The nozzle (10) is an adjustable nozzle.

7. The three-rotor multi-stage combustion turbine engine according to claim 6, characterized in that, Also includes: The mixing chamber (15) is located at the end of the afterburner (9).

8. A method for controlling a three-rotor multi-stage combustion turbine engine, used to control the three-rotor multi-stage combustion turbine engine as described in claim 7, characterized in that, include: When cruising at low Mach numbers or taking off, the three-rotor multi-stage combustion turbine engine is set to a low bypass ratio three-stage combustion mode, the second bypass valve (11) is closed, the first bypass front valve (12) is open, the first bypass rear valve (13) is closed, the rear bypass valve (14) is open, and the main combustion chamber is working. When accelerating or cruising at high Mach numbers, the three-rotor multi-stage combustion turbine is set to a low-pressure two-stage combustion mode, the second bypass valve (11) is open, the first bypass front valve (12) is closed, the first bypass rear valve (13) is open, the rear bypass valve (14) is closed, and the main combustion chamber is not working.

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