Guansheng turbine variable cycle turbine rocket engine and engine thrust realization method

Abstract

The present invention discloses a Guansheng turbine variable cycle turbo rocket engine and an engine thrust realization method. The engine is provided with two independent gas generator systems: a conventional high-pressure core engine system and a rocket gas generator system, which share a low-pressure rotor system and constitute two different working modes: a mixed exhaust turbo fan and an air turbine rocket mode; the low-pressure rotor system is provided with an integrated low-pressure rotor turbine which is located in two inner and outer ducts; the method comprises: when the engine changes from a flight Mach number below 2.5 to a Mach number above 2.5, the mixed exhaust turbo fan mode is switched to an air turbine rocket mode, and when the engine decelerates from a flight Mach number above 2.5 to a flight Mach number below 2.5, the air turbine rocket mode is switched to a mixed exhaust turbo fan mode. When the flight Mach number reaches 2.5, the present invention switches from the mixed exhaust turbo fan mode to the air turbine rocket mode, and the engine steady-state thrust increases sharply.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace vehicle propulsion, and in particular relates to a crown turbine variable cycle turbine rocket engine and an engine thrust realization method. Background Art

[0002] At present, hypersonic aircraft in the atmosphere have become a hot issue in the future aerospace field. Gas turbine engines are considered to be the optimal low-speed air-breathing power solution. Their significant specific impulse advantage can enable aircraft using gas turbine engines as a power solution to have very high operability and safety during horizontal takeoff and landing and low-speed flight.

[0003] However, the traditional gas turbine engine is limited by its own cycle heating and working characteristics. It is difficult for it to work in the speed range of Mach number 2.5 to 4.0 and generate sufficient thrust, so it is impossible for the aircraft to achieve effective flight and acceleration. There are three reasons for this: First, the airflow state at the inlet of the main combustion chamber of the existing gas turbine engine is directly related to the airflow state at the engine inlet, that is, as the engine working Mach number increases, the total temperature of the engine inlet airflow increases, which will lead to a corresponding increase in the total temperature of the main combustion chamber inlet airflow; because the total temperature at the outlet of the main combustion chamber is limited by the calorific value of the hydrocarbon fuel and the equivalence ratio of the main combustion chamber, there is an upper limit, so the higher the engine working Mach number, the smaller the main combustion chamber temperature increase for the unit airflow, the smaller the cycle heating of the engine thermodynamic cycle, and the weaker the ability to generate thrust. The second reason is that the greater the operating Mach number of a traditional gas turbine engine, the greater its inlet total temperature, and the lower the converted speed of its compression components. According to the working characteristics of general turbomachinery compression components, the lower the converted speed of the compression components, the worse its ability to allow airflow to pass through. Therefore, the converted air flow at the engine inlet will decrease with the increase of the operating Mach number. This trend will further aggravate the trend of reduced engine thrust at high operating Mach numbers. The third reason is that the increase in the operating Mach number of traditional gas turbine engines directly leads to an increase in the engine's extended total temperature, and the engine's various impeller mechanical components will operate at a lower converted speed condition; and due to constraints such as insufficient circulation heating and speed limit, it is difficult for the engine to continue to maintain a higher physical speed under high operating Mach number conditions, which will cause the converted speed of the engine's various impeller mechanical components to further decrease; according to the general characteristics of the impeller mechanical components, the efficiency of the components will decrease when the converted speed is low; the reduction in the efficiency of each impeller mechanical component will cause more energy in the engine's circulation heating to be used to heat the airflow rather than being converted into kinetic energy of the airflow, which increases the engine's heat rejection loss, reduces both thermal efficiency and overall efficiency, and further reduces thrust and specific impulse.

[0004] In summary, although the theoretical upper limit of the working Mach number of the ideal thermodynamic cycle of a traditional gas turbine engine (when both thermal efficiency and total efficiency are reduced to 0) can reach Mach 4.2, in actual engineering use, due to complex reasons such as the inability of various engine components to reach a completely ideal state, the difficulty of the engine thrust to maintain flight requirements, or the surge of the engine compression components, the upper limit of the working Mach number of a gas turbine engine in a real use environment can generally only reach between Mach 2.5 and 3. This has become a basic consensus in the field of aviation gas turbine engine research. Summary of the invention

[0005] In view of the deficiencies in the prior art, the present invention proposes a Crown turbine variable cycle turbine rocket engine and an engine thrust realization method, aiming to solve the problem that traditional gas turbine engines have insufficient cycle heating under high operating Mach number conditions and are difficult to operate in the Mach number 2.5 to 4.0 speed range and generate sufficient thrust.

[0006] The present invention proposes the following technical solutions to solve the technical problems:

[0007] A variable cycle turbo rocket engine using a crown turbine system, the engine is provided with two sets of independent gas generator systems, one is a conventional high-pressure core engine system B, and the other is a rocket gas generator system C; the two sets of independent gas generator systems share a set of low-pressure rotor system A, and share a set of branch duct flow system D and afterburning thrust-increasing system E, thereby forming two different working modes of the engine, the two different working modes are: mixed exhaust turbofan mode and air turbine rocket mode; when the conventional high-pressure core engine system B is working, the engine works in the mixed exhaust turbofan mode; when the rocket gas generator system C is working, the engine works in the air turbine rocket mode; when the flight Mach number is within the range of 0.0 to 2.5, the mixed exhaust turbofan mode is used, and when the flight Mach number is within the range of 2.5 to 4.0, the air turbine rocket mode is used; when the two sets of independent gas generator systems, one system is turned on and the other system is turned off;

[0008] Its characteristics are: the low-pressure rotor system A is provided with a low-pressure rotor turbine which is integrally manufactured but is located in two inner and outer ducts, among which the one located in the outer rocket duct D4 is the crown turbine A4, and the one located in the inner inner duct D3 is the low-pressure turbine A3; the moving blades of the crown turbine A4 are fixedly connected above the blade crowns of the moving blades of the low-pressure turbine A3; when the gas drives the low-pressure turbine A3 or the crown turbine A4, both can drive the low-pressure rotating shaft A2 and the fan A1 to rotate at the same time; when the engine works in the mixed exhaust turbofan mode, the oxygen-rich gas at the outlet of the conventional core engine system B drives the low-pressure turbine A3 to rotate the fan A1 to inhale and compress the air flow; when the engine works in the air turbine rocket mode, the fuel-rich gas at the outlet of the rocket gas generator system C drives the crown turbine A4 to rotate the fan A1 to inhale and compress the air flow; the gas flow and the air flow are mixed and re-ignited in the reburning thrust-increasing system E, and are discharged at high speed to generate thrust.

[0009] Furthermore, the low-pressure rotor system is composed of a fan A1, a low-pressure turbine A3, a crown turbine A4 and a low-pressure rotating shaft A2; wherein, the crown turbine A4 moving blades are fixedly connected above the blade crowns of the low-pressure turbine A3 moving blades, and the low-pressure turbine A3 moving blades and the crown turbine A4 moving blades rotate as a whole at the same angular velocity; there is a comb sealing structure at the edge of the blade crown of the low-pressure turbine A3 moving blades, so that the airflows in the inner duct D3 where the low-pressure turbine is located and the rocket duct D4 where the crown turbine is located are independent of each other; when the mixed exhaust turbofan mode is adopted, the oxygen-rich gas flow generated by the combustion in the main combustion chamber B4 passes through the high-pressure turbine B5 to blow the low-pressure turbine A3, and the low-pressure turbine A3 rotates while driving the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1; when the air turbine rocket mode is adopted, the fuel-rich gas flow generated by the combustion in the rocket combustion chamber C1 blows the crown turbine A4, and the crown turbine A4 rotates while driving the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1.

[0010] Furthermore, the conventional high-pressure core engine system B includes a high-pressure compressor B2, a high-pressure rotating shaft B3, a main combustion chamber B4, and a high-pressure turbine B5; the main combustion chamber B4 burns to produce oxygen-rich combustion gas to drive the high-pressure turbine B5 to rotate, and the rotation of the high-pressure turbine B5 drives the high-pressure rotating shaft B3 and the high-pressure compressor B2 to rotate together, sucking in part of the air flow passing through the fan A1 and supercharging it; the conventional high-pressure core engine system B is also provided with a front mode selection valve B1 that can be actively adjusted to open and close at the inlet of the duct, and the front mode selection valve B1 can achieve complete opening / closing control of the conventional high-pressure core engine system B and the duct D3 where the low-pressure turbine A3 is located by synchronously actuating with the rear variable area duct adjustment mechanism D5.

[0011] Furthermore, the rocket gas generator system C includes a rocket combustion chamber C1 and a rocket nozzle C2; ​​the rocket combustion chamber C1 burns to produce fuel-rich gas, and the fuel-rich gas flows into the Guansheng turbine A4 to expand and do work, and the Guansheng turbine A4 drives the low-pressure shaft to rotate A2 so that the fan A1 pressurizes the airflow sucked into the engine.

[0012] Furthermore, the sub-duct flow system D includes a bypass ring D1 before the high-pressure compressor inlet, an outer duct D2, an inner duct D3, a rocket duct D4, and a rear variable area duct regulating mechanism D5; the outer duct D2 and the inner duct D3 are formed by the bypass ring D1 before the high-pressure compressor inlet; the air flow passing through the outer duct D2 eventually flows directly into the mixer E1; the air flow passing through the inner duct D3 eventually flows into the mixer E1 via the rear variable area duct regulating mechanism D5; the rocket duct D4 is the flow channel where the fuel-rich gas flow generated by the rocket gas generator system C is located, and after flowing through the crown turbine A4, it finally flows into the mixer E1 via the rear variable area duct regulating mechanism D5.

[0013] Furthermore, the rear variable area duct regulating mechanism D5 can simultaneously control the airflow of the inner duct D3 and the rocket duct D4; when working in the mixed exhaust turbofan mode, the rear variable area duct regulating mechanism D5 adjusts and closes the rocket duct D4 where the crown turbine A4 is located; when using the air turbine rocket mode, the rear variable area duct regulating mechanism D5 cooperates with the front mode selection valve B1 to adjust and close the inner duct D3 where the conventional high-pressure core engine system B and the low-pressure turbine A3 are located.

[0014] Furthermore, the main combustion chamber B4 adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-rich combustion to ensure that when the mixed exhaust turbofan mode is used, the downstream dual-mode afterburner combustion chamber E2 has sufficient oxygen for afterburning combustion.

[0015] Furthermore, the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant pressure combustion system. In order to control the total temperature at the outlet of the rocket combustion chamber C1, the rocket combustion chamber C1 adopts rich combustion. At the same time, it can be ensured that when working in the air turbine rocket mode, the downstream dual-mode afterburner combustion chamber E2 can complete secondary combustion without the need for additional kerosene supply.

[0016] Furthermore, the afterburning thrust-increasing system E includes a mixer E1, a dual-mode afterburner combustion chamber E2, and a tail nozzle E3; when working in the mixed exhaust turbofan mode, the mixer E1 is used to mix the air flow flowing in from the outer duct D2 and the oxygen-rich fuel gas flow flowing in from the inner duct D3; when working in the air turbine rocket mode, the mixer E1 is used to mix the air flow flowing in from the outer duct D2 and the fuel-rich fuel gas flow flowing in from the rocket duct D4; the airflow mixed by the mixer E1 enters the dual-mode afterburner combustion chamber E2 for afterburning combustion; when working in the mixed exhaust turbofan mode, the dual-mode afterburner combustion chamber E2 can freely adjust whether to perform afterburning afterburning combustion by controlling the kerosene supply; when working in the air turbine rocket mode, the dual-mode afterburner combustion chamber E2 no longer supplies additional kerosene for combustion; the high-temperature and high-pressure airflow flowing out of the dual-mode afterburner combustion chamber E2 is finally discharged at high speed through the geometrically adjustable convergent-divergent structure tail nozzle E3 to generate thrust.

[0017] A method for changing the engine thrust of a variable cycle turbo rocket engine based on a crown turbine system with a flight Mach number, characterized in that:

[0018] Step 1: operate in a mixed exhaust turbofan mode within a flight Mach number range of 0.0 to 2.5;

[0019] Step 2: operate in air turbine rocket mode within the range of flight Mach number 2.5 to 4.0;

[0020] Step 3: When the engine accelerates from a flight Mach number below 2.5 to a flight Mach number above 2.5, it switches from the mixed exhaust turbofan mode to the air turborocket mode; when the engine decelerates from a flight Mach number above 2.5 to a flight Mach number below 2.5, it switches from the air turborocket mode to the mixed exhaust turbofan mode.

[0021] Furthermore, the specific process of switching from the air turbine rocket mode to the mixed exhaust turbofan mode is as follows:

[0022] 1) The engine gradually opens the front mode selection valve B1;

[0023] 2) The rear variable area duct adjustment mechanism D5 gradually expands outward;

[0024] 3) Conventional high-pressure core system B enters the windmill state, and the speed gradually increases;

[0025] 4) After the speed of the high-pressure compressor B2 reaches the critical value for ignition, the main combustion chamber B4 ignites, and at the same time, the rocket gas generator system C gradually throttles, and the low-pressure turbine A3 and the crown turbine A4 jointly drive the fan A1 to rotate;

[0026] 5) After the speed of the high-pressure compressor B2 reaches the critical value for independent operation, the rocket combustion chamber C1 is completely shut down, the low-pressure turbine A3 drives the fan A1 to rotate alone, and the dual-mode afterburner E2 is completely shut down;

[0027] 6) Continue to adjust the rear variable area duct adjustment mechanism D5 to expand outward until the rocket duct D4 is completely closed. At this point, the engine completes the mode conversion from air turbine rocket mode to mixed exhaust turbofan mode.

[0028] Furthermore, the specific process of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows:

[0029] 1) First, gradually close the engine mode selection valve B1;

[0030] 2) The rear variable area duct adjustment mechanism D5 gradually retracts to open the rocket duct D4;

[0031] 3) The rocket combustion chamber C1 begins to ignite and gradually produces rich fuel gas, the crown turbine A4 and the low-pressure turbine (A3) drive the fan A1 at the same time; the dual-mode afterburner combustion chamber E2 gradually reduces the kerosene supply to control the total residual oxygen coefficient of the engine;

[0032] 4) When the fuel-rich gas mass flow rate produced by the rocket combustion chamber C1 reaches a critical value, the kerosene supply to the main combustion chamber B4 and the dual-mode afterburner combustion chamber E2 is cut off;

[0033] 5) Continue to synchronously adjust the front mode selection valve B1 and the rear variable area duct adjustment mechanism D5 to retract until the duct D3 is completely closed. The conventional high-pressure core engine system B gradually slows down and stops working. At this point, the engine completes the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode.

[0034] Furthermore, when the engine operates in the mixed exhaust turbofan mode within the flight Mach number range of 0.0 to 2.5, the thrust and specific impulse characteristics reach the set values ​​within the Ma 0.0 to 2.5 speed domain; when the engine operates in the air turbine rocket mode within the flight Mach number range of 2.5 to 4.0, the engine steady-state thrust increases sharply, and the engine specific impulse decreases sharply due to the use of the rocket engine.

[0035] Advantages and effects of the present invention

[0036] 1. In terms of overall performance, the scheme of the present invention inherits the good thrust and specific impulse characteristics of the mixed exhaust turbofan engine in the speed range of Ma 0.0 to 2.5. At the same time, the thrust advantage of the air turbine rocket in the speed range of Mach number 2.5 to 4.0 makes up for the defects of the traditional gas turbine engine in this speed range that the circulation heating is insufficient, the thrust cannot be generated, or even the engine cannot work: because the rocket combustion chamber combustion and energy addition of the air turbine rocket are no longer constrained by the state of the engine inlet airflow, the core technical bottleneck of the conventional gas turbine engine in the speed range of Mach number 2.5 to 4.0, which has a significant lack of thrust, a significant reduction in fuel efficiency, or even an inability to work normally, can be effectively solved by converting the engine mode to the air turbine rocket mode. Provide power, thus constructing a breath-in variable cycle engine with feasible principles and high engineering feasibility.

[0037] 2. The present invention adopts a crown turbine structure, which reduces the difficulty of structural implementation of parallel design of two different gas generators, maintains the compact axial layout of the engine, and brings positive benefits to shortening the rotating shaft and improving the shaft strength; at the same time, it provides a more significant process control benefit for the two-way mode conversion process between the dual working modes of the engine, so that the start-up and shutdown processes of the two core engines will not cause the low-pressure shaft rotor to stop, providing a feasibility basis for smooth thrust mode conversion.

[0038] 3. The present invention shares the low-pressure turbine scheme and the crown turbine scheme. The thrust in the air-turbine rocket mode increases with the increase of the flight Mach number, which can better meet the power requirements of the aircraft under high-speed flight conditions; at the same time, it is also more conducive to the thrust connection with the air-breathing power scheme with a higher Mach number, and constructs a combined cycle engine scheme that meets the requirements of wide-speed hypersonic flight. When the flight Mach number reaches 2.5, the scheme of the present invention switches from the mixed exhaust turbofan mode to the air-turbine rocket mode. Although the engine specific impulse decreases sharply due to the use of the rocket engine, the engine steady-state thrust increases sharply. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1-1 It is a vertical cross-sectional view of a variable cycle turbine rocket engine of a Guansheng turbine system of the present invention;

[0040] Figure 1-2 One of the partial vertical cross-sections of the variable cycle turbo rocket engine that invented the Guansheng turbine system;

[0041] Figure 1-3 The second partial vertical cross-sectional view of the variable cycle turbine rocket engine of the Guansheng turbine system of the present invention;

[0042] Figure 2-1 A schematic diagram of a mixed exhaust turbofan mode of the Guansheng turbine system of the present invention;

[0043] Figure 2-2 It is a schematic diagram of the air turbine rocket mode of the Guansheng turbine system of the present invention;

[0044] Figure 3 This is a schematic diagram of how the system thrust changes with the flight Mach number;

[0045] Figure 4 This is a schematic diagram of the variation of system specific impulse with flight Mach number;

[0046] Figure 5 is the variation of the system inlet air flow rate with the flight Mach number;

[0047] Figure 6 The available pressure ratio of the system tail nozzle varies with the flight Mach number;

[0048] Figure 7 The total temperature at the inlet of the system's low-pressure turbine varies with the flight Mach number. DETAILED DESCRIPTION

[0049] Design principle of the present invention

[0050] 1. Reasons for adding the engine of the present invention to the rocket combustion chamber: The working state of the original gas turbine engine main combustion chamber is directly constrained by the engine inlet conditions, and the engine cycle heating is directly limited by the cycle compression ratio and fuel usage. The present invention adds an independent rocket combustion chamber on the basis of the original main combustion chamber. Since the rocket combustion chamber does not use the air entering the engine as an oxidant, its combustion state and cycle heating are no longer affected by the changes in the engine inlet airflow state, that is, the rocket combustion chamber combustion state and the engine operating Mach number no longer have a direct constraint relationship. The decoupling of this strong constraint relationship solves the core technical bottleneck of insufficient cycle heating of traditional gas turbine engines under high operating Mach number conditions, and provides a feasible technical path for achieving stable engine operation and generating sufficient thrust.

[0051] The reason why the combustion state of the rocket combustion chamber is not affected by air is that it controls the mass mixing ratio of the oxidant and the reductant of the liquid oxygen-kerosene combustion system in the rocket combustion chamber C1, so that the total inlet temperature of the common low-pressure turbine A3 is maintained at 1200-1500K, so as to ensure that the total inlet temperature of the common low-pressure turbine A3 in the two working modes is adapted to each other; it can be seen that the combustion state of its combustion chamber is only related to the mass mixing ratio of the oxidant and the reductant of the liquid oxygen-kerosene combustion system, and has nothing to do with air. The reason why the combustion state of the main combustion chamber of the gas turbine combustion chamber is related to air is that the main combustion chamber B4 adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-enriched combustion.

[0052] 2. Difficulties in design of the present invention:

[0053] Since the engine adds a rocket combustion chamber on the basis of the traditional combustion chamber, there are two design schemes for the two combustion chambers and their supporting impeller machinery systems, namely, a parallel structural layout or a series structural layout. The parallel structural layout means that the main combustion chamber and its supporting impeller machinery, and the rocket combustion chamber and its supporting impeller machinery are located at two different rotation axes, and the two rotation axes are arranged in parallel. Each scheme of the present invention adopts a design method for the series structural layout of two combustion chambers and their supporting impeller machinery. The series structural layout means that the main combustion chamber and its supporting impeller machinery, and the rocket combustion chamber and its supporting impeller machinery are on the same rotation axis.

[0054] a. The parallel structure layout encounters the difficulty of "dead weight". The aviation gas turbine engine and air turbine rocket engine with a parallel structure layout bring great difficulties to the combined design of the intake and exhaust systems used by the matching engines, and it is difficult to solve the significant "dead weight" problem caused when any engine is not working in different working speed domains;

[0055] b. The difficulty of inconsistent thermal cycle design requirements encountered in the series structure layout. Aviation gas turbine engines and air turbine rocket engines with a series structure layout need to solve the problem of inconsistent requirements for key thermal cycle design parameters of the engine bypass ratio and pressure ratio within their respective advantageous operating speed ranges;

[0056] For the mixed exhaust turbofan mode, in order to enable this working mode to accelerate to Mach 2.5 and to maximize the engine's performance between Mach 0 and 2.5, the engine needs to make a reasonable compromise design for its thermodynamic cycle design parameters. The engine will adopt a total pressure ratio of about 10 to 25 and a total bypass ratio of about 0.5 to 1.5, so that the engine can generate sufficient thrust between Mach numbers 0 and 2.5, and achieve a high specific impulse to reduce fuel consumption during flight acceleration; at the same time, the engine shaft speed limit, the temperature limit of the hot end components, and the aerodynamic stability limit of the compression components can also be met.

[0057] For the air turbine rocket mode, in order to make the engine produce enough thrust while maintaining a specific impulse that is not too low, the selection of the engine's thermal cycle design parameters is significantly different from that of the mixed exhaust turbofan mode. The use of a total bypass ratio design of 5 to 8 can, on the one hand, improve the engine's total residual oxygen coefficient, increase the temperature of secondary combustion and thus increase thrust; on the other hand, it can effectively increase the engine's specific impulse and minimize fuel consumption during flight acceleration. The use of a total boost ratio design of 3 to 6 can, on the one hand, reduce the demand for turbine output power, so that the engine can use fewer stages of turbines to drive the fan to compress and work, effectively reducing the weight of the engine structure; on the other hand, it can reduce the combustion pressure of the rocket combustion chamber and reduce the design difficulty of the rocket combustion chamber and rocket propellant supply system.

[0058] In short, there are significant differences in the boost capacity and bypass ratio requirements of the compressors of the two modes of engines, and this problem needs to be solved when adopting a series structure layout.

[0059] 2. Solution of the present invention

[0060] The present invention aims at solving the problem of inconsistent requirements on the boosting capacity and bypass ratio of the compressor of two-mode engines, and adopts an inner and outer bypass design method that can take into account the requirements of the two modes at the same time.

[0061] 1) For the mixed exhaust turbofan mode of the Guansheng turbine system, the present invention adopts the inner duct to synchronize the front mode selection valve B1 and the rear variable area duct adjustment mechanism D5 to realize the complete opening / closing control of the inner duct D3 where the conventional high-pressure core engine system B and the low-pressure turbine A3 are located. In this way, the high-pressure compressor B2 in the inner duct D3 can pressurize the airflow again, so that the total boost ratio of the engine is higher in the mixed exhaust turbofan mode. At the same time, the high-pressure compressor B2 has a better air intake capacity, so that more airflow flows into the inner duct D3, and the total bypass ratio of the engine is smaller, which is more in line with the trend of better engine performance in the mixed exhaust turbofan mode;

[0062] 2) In the air turbine rocket mode of the Guansheng turbine system of the present invention, the rear variable area duct adjustment mechanism D5 cooperates with the front mode selection valve B1 to adjust and close the inner duct D3 where the conventional high-pressure core engine system B and the low-pressure turbine A3 are located. The airflow is forced to flow into the outer duct D2 at D1, and the total bypass ratio of the engine reaches a relatively large state by controlling the rich fuel gas flow of the rocket duct D4. At the same time, since the inner duct D3 is closed, the high-pressure compressor B2 stops working, and the boost capacity of the fan A1 is low, so the demand for a low total boost ratio can be met.

[0063] 3) For the air turbine rocket mode of the crown turbine system, the present invention adopts the method of crown turbine system A4 on the low-pressure turbine A3 blade crown, which reduces the difficulty of parallel design of two different gas generators, maintains the compact axial layout of the engine, and brings positive benefits to shortening the rotating shaft and improving the shaft strength; at the same time, it provides a more significant process control benefit for the two-way mode conversion process between the dual working modes of the engine, so that the startup and shutdown processes of the two core engines will not cause the low-pressure shaft rotor to stop, providing a feasible basis for the smooth thrust mode conversion.

[0064] The present invention is further explained below in conjunction with the accompanying drawings:

[0065] A variable cycle turbo rocket engine using a crown turbine system Figure 1-1 , 1-2 As shown in 1-3, the engine is provided with two sets of independent gas generator systems, one set is a conventional high-pressure core engine system B, and the other set is a rocket gas generator system C; the two sets of independent gas generator systems share a set of low-pressure rotor system A, and share a set of sub-duct flow system D and afterburning thrust system E, thereby forming two different working modes of the engine, and the two different working modes are: mixed exhaust turbofan mode and air turbine rocket mode; when the conventional high-pressure core engine system B is working, the engine works in the mixed exhaust turbofan mode; when the rocket gas generator system C is working, the engine works in the air turbine rocket mode; when the flight Mach number is within the range of 0.0 to 2.5, the mixed exhaust turbofan mode is used, and when the flight Mach number is within the range of 2.5 to 4.0, the air turbine rocket mode is used; the two sets of independent gas generator systems, when one system is turned on, the other system is turned off;

[0066] Its characteristics are: Figure 1-2 , 1-3 As shown, the low-pressure rotor system A is provided with a low-pressure rotor turbine which is integrally manufactured but is located in two inner and outer ducts, wherein the one located in the outer rocket duct D4 is the crown turbine A4, and the one located in the inner inner duct D3 is the low-pressure turbine A3; the moving blades of the crown turbine A4 are fixedly connected above the blade crowns of the moving blades of the low-pressure turbine A3; when the gas drives the low-pressure turbine A3 or the crown turbine A4, both can drive the low-pressure rotating shaft A2 and the fan A1 to rotate at the same time; when the engine works in the mixed exhaust turbofan mode, the oxygen-rich gas at the outlet of the conventional core engine system B drives the low-pressure turbine A3 to rotate the fan A1 to inhale and compress the air flow; when the engine works in the air turbine rocket mode, the fuel-rich gas at the outlet of the rocket gas generator system C drives the crown turbine A4 to rotate the fan A1 to inhale and compress the air flow; the gas flow and the air flow are mixed and re-ignited in the re-combustion thrust-increasing system E, and are discharged at high speed to generate thrust.

[0067] Supplementary Note 1

[0068] Guansheng turbine solution Figure 2-1 , 2-2 As shown, when the mixed exhaust turbofan mode is adopted, Figure 2-1 , the rocket combustion chamber C1 of the rocket gas generator system C is closed, and the main combustion chamber B4 of the conventional high-pressure core engine system B is opened; when the air turbine rocket mode is adopted, such as Figure 2-2 , the main combustion chamber B4 of the conventional high-pressure core engine system B is closed, and the rocket combustion chamber C1 of the rocket gas generator system C is opened.

[0069] Furthermore, if Figure 1-2 As shown, the low-pressure rotor system consists of a fan A1, a low-pressure turbine A3, a crown turbine A4 and a low-pressure rotating shaft A2; wherein, the crown turbine A4 moving blades are fixedly connected to the top of the low-pressure turbine A3 moving blades, and the low-pressure turbine A3 moving blades and the crown turbine A4 moving blades rotate as a whole at the same angular velocity; there is a comb seal structure at the edge of the low-pressure turbine A3 moving blade crown, so that the airflows in the inner duct D3 where the low-pressure turbine is located and the rocket duct D4 where the crown turbine is located are independent of each other; when the mixed exhaust turbofan mode is adopted, the oxygen-rich gas flow generated by the combustion in the main combustion chamber B4 passes through the high-pressure turbine B5 to blow the low-pressure turbine A3, and the low-pressure turbine A3 rotates while driving the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1; when the air turbine rocket mode is adopted, the fuel-rich gas flow generated by the combustion in the rocket combustion chamber C1 blows the crown turbine A4, and the crown turbine A4 rotates while driving the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1.

[0070] Furthermore, if Figure 1-2 As shown, the conventional high-pressure core engine system B includes a high-pressure compressor B2, a high-pressure rotating shaft B3, a main combustion chamber B4, and a high-pressure turbine B5; the main combustion chamber B4 burns to produce oxygen-rich combustion gas to drive the high-pressure turbine B5 to rotate, and the rotation of the high-pressure turbine B5 drives the high-pressure rotating shaft B3 and the high-pressure compressor B2 to rotate together, sucking in part of the air flow passing through the fan A1 and supercharging it; the conventional high-pressure core engine system B is also provided with a front mode selection valve B1 that can be actively adjusted to open and close at the inlet of the duct, and the front mode selection valve B1 can achieve full opening / closing control of the conventional high-pressure core engine system B and the duct D3 where the low-pressure turbine A3 is located by synchronously actuating with the rear variable area duct adjustment mechanism D5.

[0071] Furthermore, if Figure 1-2As shown, the rocket gas generator system C includes a rocket combustion chamber C1 and a rocket nozzle C2; ​​the rocket combustion chamber C1 burns to produce fuel-rich gas, and the fuel-rich gas flows into the Guansheng turbine A4 to expand and do work, and the Guansheng turbine A4 drives the low-pressure shaft to rotate A2 so that the fan A1 pressurizes the airflow sucked into the engine.

[0072] Furthermore, if Figure 1-3 As shown, the sub-duct flow system D includes a bypass ring D1 before the high-pressure compressor inlet, an outer duct D2, an inner duct D3, a rocket duct D4, and a rear variable area duct regulating mechanism D5; the outer duct D2 and the inner duct D3 are formed by diverting at the bypass ring D1 before the high-pressure compressor inlet; the air flow passing through the outer duct D2 eventually flows directly into the mixer E1; the air flow passing through the inner duct D3 eventually flows into the mixer E1 via the rear variable area duct regulating mechanism D5; the rocket duct D4 is the flow channel where the fuel-rich gas flow generated by the rocket gas generator system C is located, and after flowing through the crown turbine A4, it finally flows into the mixer E1 via the rear variable area duct regulating mechanism D5.

[0073] Furthermore, if Figure 1-3 As shown, the rear variable area duct regulating mechanism D5 can simultaneously control the airflow of the inner duct D3 and the rocket duct D4; when using the mixed exhaust turbofan mode, the rear variable area duct regulating mechanism D5 adjusts and closes the rocket duct D4 where the crown turbine A4 is located; when using the air turbine rocket mode, the rear variable area duct regulating mechanism D5 cooperates with the front mode selection valve B1 to adjust and close the inner duct D3 where the conventional high-pressure core engine system B and the low-pressure turbine A3 are located.

[0074] Furthermore, if Figure 1-2 As shown, the main combustion chamber B4 adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-rich combustion to ensure that when the mixed exhaust turbofan mode is used, the downstream dual-mode afterburner combustion chamber E2 has sufficient oxygen for afterburning and afterburning.

[0075] Furthermore, the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant pressure combustion system. In order to control the total temperature at the outlet of the rocket combustion chamber C1, the rocket combustion chamber C1 adopts rich combustion. At the same time, it can be ensured that when working in the air turbine rocket mode, the downstream dual-mode afterburner combustion chamber E2 can complete secondary combustion without the need for additional kerosene supply.

[0076] Furthermore, if Figure 1-3As shown, the afterburning thrust-increasing system E includes a mixer E1, a dual-mode afterburner combustion chamber E2, and a tail nozzle E3; when working in the mixed exhaust turbofan mode, the mixer E1 is used to mix the air flow flowing in from the outer duct D2 and the oxygen-rich fuel gas flow flowing in from the inner duct D3; when working in the air turbine rocket mode, the mixer E1 is used to mix the air flow flowing in from the outer duct D2 and the fuel-rich fuel gas flow flowing in from the rocket duct D4; the airflow mixed by the mixer E1 enters the dual-mode afterburner combustion chamber E2 for afterburning combustion; when working in the mixed exhaust turbofan mode, the dual-mode afterburner combustion chamber E2 can freely adjust whether to perform afterburning afterburning combustion by controlling the kerosene supply; when working in the air turbine rocket mode, the dual-mode afterburner combustion chamber E2 no longer supplies additional kerosene for combustion; the high-temperature and high-pressure airflow flowing out of the dual-mode afterburner combustion chamber E2 is finally discharged at high speed through the geometrically adjustable convergent-divergent structure tail nozzle E3 to generate thrust.

[0077] A method for varying engine thrust with flight Mach number based on a crown turbine, characterized in that:

[0078] Step 1: operate in a mixed exhaust turbofan mode within a flight Mach number range of 0.0 to 2.5;

[0079] Step 2: operate in air turbine rocket mode within the range of flight Mach number 2.5 to 4.0;

[0080] Step 3: When the engine accelerates from a flight Mach number below 2.5 to a flight Mach number above 2.5, it switches from the mixed exhaust turbofan mode to the air turborocket mode; when the engine decelerates from a flight Mach number above 2.5 to a flight Mach number below 2.5, it switches from the air turborocket mode to the mixed exhaust turbofan mode.

[0081] Furthermore, the specific process of switching from the air turbine rocket mode to the mixed exhaust turbofan mode is as follows:

[0082] 1) The engine gradually opens the front mode selection valve B1;

[0083] 2) The rear variable area duct adjustment mechanism D5 gradually expands outward;

[0084] 3) Conventional high-pressure core system B enters the windmill state, and the speed gradually increases;

[0085] 4) After the speed of the high-pressure compressor B2 reaches the critical value for ignition, the main combustion chamber B4 ignites, and at the same time, the rocket gas generator system C gradually throttles, and the low-pressure turbine A3 and the crown turbine A4 jointly drive the fan A1 to rotate;

[0086] 5) After the speed of the high-pressure compressor B2 reaches the critical value for independent operation, the rocket combustion chamber C1 is completely shut down, the low-pressure turbine A3 drives the fan A1 to rotate alone, and the dual-mode afterburner E2 is completely shut down;

[0087] 6) Continue to adjust the rear variable area duct adjustment mechanism D5 to expand outward until the rocket duct D4 is completely closed. At this point, the engine completes the mode conversion from air turbine rocket mode to mixed exhaust turbofan mode.

[0088] Furthermore, the specific process of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows:

[0089] 1) First, gradually close the engine mode selection valve B1;

[0090] 2) The rear variable area duct adjustment mechanism D5 gradually retracts to open the rocket duct D4;

[0091] 3) The rocket combustion chamber C1 begins to ignite and gradually produces rich fuel gas, the crown turbine A4 and the low-pressure turbine (A3) drive the fan A1 at the same time; the dual-mode afterburner combustion chamber E2 gradually reduces the kerosene supply to control the total residual oxygen coefficient of the engine;

[0092] 4) When the fuel-rich gas mass flow rate produced by the rocket combustion chamber C1 reaches a critical value, the kerosene supply to the main combustion chamber B4 and the dual-mode afterburner combustion chamber E2 is cut off;

[0093] 5) Continue to synchronously adjust the front mode selection valve B1 and the rear variable area duct adjustment mechanism D5 to retract until the duct D3 is completely closed. The conventional high-pressure core engine system B gradually slows down and stops working. At this point, the engine completes the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode.

[0094] Furthermore, when the engine operates in the mixed exhaust turbofan mode within the flight Mach number range of 0.0 to 2.5, the thrust and specific impulse characteristics reach the set values ​​within the speed range of Mach numbers 0.0 to 2.5; when the engine operates in the air turbine rocket mode within the flight Mach number range of 2.5 to 4.0, the engine steady-state thrust increases sharply, and the engine specific impulse decreases sharply due to the use of the rocket engine.

[0095] Example 1: Experimental results of two modes

[0096] The present invention adds the effect of increasing thrust in the rocket combustion chamber. Figure 3 , Figure 4 As shown, Figure 3 It shows that when the Mach number is greater than 2.5, the rocket's thrust increases significantly. Figure 4It shows that when the Mach number is greater than 2.5, the specific impulse of the rocket decreases. From the comparison of these two pictures, when the Mach number is greater than 2.5, although the specific impulse of the rocket combustion chamber decreases, that is, the unit thrust of the fuel decreases, or the fuel consumed for the same thrust increases, the thrust of the rocket increases significantly. When the Mach number is less than 2.5, the mixed exhaust turbofan mode is adopted. At this time, the thrust generated by the combustion chamber of the B system gradually decreases as the Mach number increases, and shows a rapid downward trend when it is about to reach Mach number 2.5. When the Mach number is greater than 2.5, the air turbine rocket mode based on the rocket combustion chamber is adopted. At this time, the thrust generated by the C system gradually increases and does not weaken due to the increase in the Mach number.

[0097] The invention adds a rocket combustion chamber, so that the inlet flow rate increases. Figure 5 As shown, when the Mach number is less than 2.5, due to the use of traditional gas turbine combustion chambers, the higher the temperature of the main combustion chamber, the lower the converted speed. The lower the converted speed, the worse its ability to allow airflow to pass, that is, the lower the inlet flow rate. When the Mach number is greater than 2.5, since the inlet flow of the rocket combustion chamber is not affected by the airflow, but is only related to the combustion material, when the Mach number is greater than 2.5, the inlet flow rate shows an upward trend. As the inlet flow rate increases, the engine's operating efficiency increases, the engine speed increases, and the thrust increases. The effect of increased thrust is as follows: Figure 3 shown.

[0098] The invention adds the effect that the total temperature of the rocket combustion chamber low-pressure turbine inlet does not increase with the Mach number. Figure 7 As shown, when the Mach number is less than 2.5, the total temperature at the low-pressure turbine inlet of system B increases in proportion to the Mach number. When the Mach number is greater than 2.5, the total temperature at the low-pressure turbine inlet of system C remains constant and is not proportional to the Mach number.

[0099] It should be emphasized that the above specific embodiments are merely explanations of the present invention, and they are not limitations of the present invention. After reading this specification, those skilled in the art can make modifications to the above embodiments without any creative contribution as needed, but as long as they are within the scope of the claims of the present invention, they are protected by patent law.

Claims

1. A variable cycle turbo rocket engine using a Guansheng turbine system, the engine is provided with two sets of mutually independent gas generator systems, one is a conventional high pressure core engine system (B), and the other is a rocket gas generator system (C); characterized in that: The conventional high-pressure core engine system (B) comprises a high-pressure compressor (B2), a high-pressure rotating shaft (B3), a main combustion chamber (B4), and a high-pressure turbine (B5); the main combustion chamber (B4) generates oxygen-rich combustion gas to drive the high-pressure turbine (B5) to rotate, and the rotation of the high-pressure turbine (B5) drives the high-pressure rotating shaft (B3) and the high-pressure compressor (B2) to rotate together, sucking in part of the air flow passing through the fan (A1) and supercharging it; The conventional high-pressure core engine system (B) is also provided with a front mode selection valve (B1) that can be actively adjusted to open and close at the inlet of the duct, and the front mode selection valve (B1) can achieve complete opening / closing control of the conventional high-pressure core engine system (B) and the duct (D3) where the low-pressure turbine (A3) is located by synchronously moving with the rear variable area duct adjustment mechanism (D5); The two independent gas generator systems share a low-pressure rotor system (A), a sub-duct flow system (D) and a re-ignition thrust-increasing system (E), thereby forming two different working modes of the engine, the two different working modes are: a mixed exhaust turbofan mode and an air turbine rocket mode; when the conventional high-pressure core engine system (B) is working, the engine works in the mixed exhaust turbofan mode; when the rocket gas generator system (C) is working, the engine works in the air turbine rocket mode; when the flight Mach number is within the range of 0.0 to 2.5, the mixed exhaust turbofan mode is used, and when the flight Mach number is within the range of 2.5 to 4.0, the air turbine rocket mode is used; The two independent gas generator systems are designed so that when one system is turned on, the other system is turned off; The low-pressure rotor system (A) is provided with a low-pressure rotor turbine which is integrally manufactured but is located in two inner and outer ducts, wherein the one located in the outer rocket duct (D4) is a crown turbine (A4), and the one located in the inner inner duct (D3) is a low-pressure turbine (A3); the moving blades of the crown turbine (A4) are fixedly connected to the top of the blade crown of the moving blades of the low-pressure turbine (A3); when the low-pressure turbine (A3) or the crown turbine (A4) is driven by the gas, the low-pressure rotating shaft (A2) and the fan (A1) can be driven to rotate at the same time; when the engine works in the mixed exhaust turbofan mode, the oxygen-rich gas at the outlet of the conventional high-pressure core engine system (B) drives the low-pressure turbine (A3) to rotate the fan (A1) to inhale and compress the air flow; when the engine works in the air turbine rocket mode, the fuel-rich gas at the outlet of the rocket gas generator system (C) drives the crown turbine (A4) to rotate the fan (A1) to inhale and compress the air flow; the gas flow and the air flow are mixed and ignited for the second time in the afterburning thrust-increasing system (E), and discharged at high speed to generate thrust.

2. A variable cycle turbo rocket engine using a crown turbine system according to claim 1, characterized in that: The low-pressure rotor system is composed of a fan (A1), a low-pressure turbine (A3), a crown turbine (A4) and a low-pressure rotating shaft (A2); wherein, The moving blades of the crown turbine (A4) are fixedly connected to the top of the blade crown of the moving blades of the low-pressure turbine (A3), and the moving blades of the low-pressure turbine (A3) and the moving blades of the crown turbine (A4) rotate together at a constant angular velocity as a whole; there is a comb seal structure at the edge of the blade crown of the moving blade of the low-pressure turbine (A3), so that the airflow of the inner duct (D3) where the low-pressure turbine is located and the rocket duct (D4) where the crown turbine is located are independent of each other; when the mixed exhaust turbofan mode is adopted, the oxygen-rich gas flow generated by the combustion of the main combustion chamber (B4) passes through the high-pressure turbine (B5) to blow the low-pressure turbine (A3), and the low-pressure turbine (A3) rotates while driving the crown turbine (A4), the low-pressure rotating shaft (A2) and the fan (A1); when the air turbine rocket mode is adopted, the fuel-rich gas flow generated by the combustion of the rocket combustion chamber (C1) blows the crown turbine (A4), and the crown turbine (A4) rotates while driving the low-pressure turbine (A3), the low-pressure rotating shaft (A2) and the fan (A1).

3. A variable cycle turbo rocket engine using a crown turbine system according to claim 1, characterized in that: The rocket gas generator system (C) includes a rocket combustion chamber (C1) and a rocket nozzle (C2); the rocket combustion chamber (C1) burns to produce fuel-rich gas, and the fuel-rich gas flows into the crown turbine (A4) to expand and do work, and the crown turbine (A4) drives the low-pressure rotating shaft (A2) so that the fan (A1) pressurizes the airflow sucked into the engine.

4. A variable cycle turbo rocket engine using a crown turbine system according to claim 1, characterized in that: The sub-duct flow system (D) comprises a flow divider ring (D1) before the high-pressure compressor inlet, an outer duct (D2), an inner duct (D3), a rocket duct (D4), and a rear variable area duct regulating mechanism (D5); the outer duct (D2) and the inner duct (D3) are formed by flow division at the flow divider ring (D1) before the high-pressure compressor inlet; the air flow passing through the outer duct (D2) eventually flows directly into the mixer (E1); the air flow passing through the inner duct (D3) eventually flows into the mixer (E1) via the rear variable area duct regulating mechanism (D5); the rocket duct (D4) is the flow channel where the fuel-rich gas flow generated by the rocket gas generator system (C) is located, and after flowing through the crown turbine (A4), it finally flows into the mixer (E1) via the rear variable area duct regulating mechanism (D5).

5. A variable cycle turbo rocket engine using a crown turbine system according to claim 4, characterized in that: The rear variable area duct regulating mechanism (D5) can simultaneously control the airflow of the inner duct (D3) and the rocket duct (D4); when the mixed exhaust turbofan mode is used, the rear variable area duct regulating mechanism (D5) regulates and closes the rocket duct (D4) where the crown turbine (A4) is located; when the air turbine rocket mode is used, the rear variable area duct regulating mechanism (D5) cooperates with the front mode selection valve (B1) to regulate and close the inner duct (D3) where the conventional high-pressure core engine system (B) and the low-pressure turbine (A3) are located.

6. A variable cycle turbo rocket engine using a crown turbine system according to claim 3, characterized in that: The main combustion chamber (B4) adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-rich combustion to ensure that when the mixed exhaust turbofan mode is used, the downstream dual-mode afterburner combustion chamber (E2) has sufficient oxygen for afterburning and afterburning combustion.

7. A variable cycle turbo rocket engine using a crown turbine system according to claim 3, characterized in that: The rocket combustion chamber (C1) adopts a liquid oxygen-kerosene constant pressure combustion system. In order to control the total temperature at the outlet of the rocket combustion chamber (C1), the rocket combustion chamber (C1) adopts rich combustion. At the same time, it can be ensured that when working in the air turbine rocket mode, the downstream dual-mode afterburner combustion chamber (E2) can complete secondary combustion without the need for additional kerosene supply.

8. The variable cycle turbine rocket engine using a crown turbine system according to claim 1, characterized in that: The afterburning thrust increasing system (E) comprises a mixer (E1), a dual-mode afterburner (E2), and a tail nozzle (E3); when the mixed exhaust turbofan mode is used, the mixer (E1) is used to mix the air flow flowing in from the outer duct (D2) and the oxygen-rich gas flow flowing in from the inner duct (D3); when the air turbine rocket mode is used, the mixer (E1) is used to mix the air flow flowing in from the outer duct (D2) and the fuel-rich gas flow flowing in from the rocket duct (D4); the air flow is mixed via the mixer (E1). The afterburning airflow enters the dual-mode afterburner combustion chamber (E2) for afterburning combustion; when the mixed exhaust turbofan mode is working, the dual-mode afterburner combustion chamber (E2) can freely adjust whether to perform afterburning combustion by controlling the kerosene supply; when the air turbine rocket mode is working, the dual-mode afterburner combustion chamber (E2) no longer supplies additional kerosene for combustion; the high-temperature and high-pressure airflow flowing out of the dual-mode afterburner combustion chamber (E2) is finally discharged at high speed through a geometrically adjustable convergent-divergent tail nozzle (E3) to generate thrust.

9. A method for changing the engine thrust of a variable cycle turbo rocket engine using a crown turbine system according to any one of claims 1 to 8 with the flight Mach number, characterized in that: Step 1: operate in mixed exhaust turbofan mode within the flight Mach number range of 0.0 to 2.5; Step 2: Use the air turbine rocket mode to operate within the flight Mach number range of 2.5 to 4.0; Step 3: When the engine accelerates from a flight Mach number below 2.5 to a flight Mach number above 2.5, it switches from the mixed exhaust turbofan mode to the air turborocket mode; when the engine decelerates from a flight Mach number above 2.5 to a flight Mach number below 2.5, it switches from the air turborocket mode to the mixed exhaust turbofan mode.

10. A method for changing the thrust of an engine based on a crown turbine with the flight Mach number according to claim 9, characterized in that: The specific process of switching from the air turbine rocket mode to the mixed exhaust turbofan mode is as follows: 1) The engine gradually opens the front mode selection valve (B1); 2) The rear variable area duct adjustment mechanism (D5) gradually expands outward; 3) The conventional high-pressure core system (B) enters the windmill state, and the speed gradually increases; 4) After the speed of the high-pressure compressor (B2) reaches the critical value for ignition, the main combustion chamber (B4) ignites, and at the same time, the rocket gas generator system (C) gradually throttles, and the low-pressure turbine (A3) and the crown turbine (A4) jointly drive the fan (A1) to rotate; 5) After the speed of the high-pressure compressor (B2) reaches the critical value for independent operation, the rocket combustion chamber (C1) is completely shut down, the low-pressure turbine (A3) drives the fan (A1) to rotate independently, and the dual-mode afterburner combustion chamber (E2) is completely shut down; 6) Continue to adjust the rear variable area duct adjustment mechanism (D5) to expand outward until the rocket duct (D4) is completely closed. At this point, the engine completes the mode conversion from air turbine rocket mode to mixed exhaust turbofan mode.

11. A method for changing the thrust of an engine based on a crown turbine with the flight Mach number according to claim 10, characterized in that: The specific process of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows: 1) First, gradually close the engine mode selection valve (B1); 2) The rear variable area duct adjustment mechanism (D5) gradually retracts to open the rocket duct (D4); 3) The rocket combustion chamber (C1) begins to ignite and gradually produces rich combustion gas, the crown turbine (A4) and the low-pressure turbine (A3) simultaneously drive the fan (A1); the dual-mode afterburner (E2) gradually reduces the kerosene supply to control the total residual oxygen coefficient of the engine; 4) After the fuel-rich gas mass flow rate generated by the rocket combustion chamber (C1) reaches a critical value, the kerosene supply to the main combustion chamber (B4) and the dual-mode afterburner combustion chamber (E2) is cut off; 5) Continue to synchronously adjust the front mode selection valve (B1) and the rear variable area duct adjustment mechanism (D5) inward until the duct (D3) is completely closed. The conventional high-pressure core engine system (B) gradually slows down and stops working. At this point, the engine completes the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode.

12. The method for changing the thrust of an engine based on a crown turbine with the flight Mach number according to claim 10, characterized in that: When the engine operates in the mixed exhaust turbofan mode within the flight Mach number range of 0.0~2.5, the thrust and specific impulse characteristics reach the set values ​​in the Ma 0.0~2.5 speed domain; when the engine operates in the air turbine rocket mode within the flight Mach number range of 2.5~4.0, the engine steady-state thrust increases sharply, and the engine specific impulse decreases sharply due to the use of the rocket engine.

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