A dual-mode variable cycle turbo rocket engine

Abstract

The present invention discloses a dual-mode variable cycle turbine rocket engine, which is provided with two sets of independent gas generator systems: a conventional high-pressure core engine system and a rocket gas generator system; they share a set of low-pressure rotor system; the shared low-pressure rotor system includes a shared low-pressure turbine scheme and a crown turbine scheme; when the shared low-pressure turbine scheme is adopted, the shared low-pressure turbine is driven to rotate by the respective gases of the two gas generators in two different working modes; when the crown turbine scheme is adopted, the gas flow generated by the main combustion chamber of the conventional high-pressure core engine system blows the low-pressure turbine to rotate, and the gas flow generated by the rocket combustion chamber of the rocket gas generator system blows the crown turbine to rotate; the shared low-pressure turbine scheme and the crown turbine scheme of the present invention, when the flight Mach number reaches 2.5, the mixed exhaust turbofan mode is switched to the air turbine rocket mode, and the steady-state thrust of the engine increases sharply.

Description

Technical Field

[0001] The invention belongs to the field of aerospace vehicle propulsion, and in particular relates to a dual-mode variable cycle turbine rocket engine. 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 2.5~4.0 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, 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 dual-mode variable cycle turbine rocket engine, which aims 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 speed range of Mach numbers 2.5 to 4.0 and generate sufficient thrust.

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

[0007] A dual-mode variable cycle turbo rocket engine, characterized in that: 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, 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;

[0008] The shared low-pressure rotor system A includes a shared low-pressure turbine solution for reducing the radial size of the engine and the weight of the low-pressure turbine, and a crown turbine solution for decoupling the aerodynamic design of the low-pressure turbine;

[0009] When the shared low-pressure turbine scheme of the low-pressure rotor system A is adopted, the conventional high-pressure core engine system B and the rocket gas generator system C share the low-pressure turbine A3 of the low-pressure rotor system A. The shared low-pressure turbine A3 is driven to rotate by the respective combustion gases of the two sets of gas generators in two different working modes, and then drives the low-pressure rotating shaft A2 and the fan A1 in turn, and then the airflow is sucked into the engine by the fan A1 and compressed;

[0010] When the crown turbine scheme of the low-pressure rotor system A is adopted, the gas flow generated by the main combustion chamber B4 of the conventional high-pressure core engine system B blows the low-pressure turbine A3 to rotate, and the rotation of the low-pressure turbine A3 will simultaneously drive the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1 to rotate; the gas flow generated by the rocket combustion chamber C1 of the rocket gas generator system C blows the crown turbine A4 to rotate, and the rotation of the crown turbine A4 will simultaneously drive the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1 to rotate.

[0011] Furthermore, when a common low-pressure turbine scheme of the low-pressure rotor system A is adopted, the low-pressure rotor system A is provided with a fan A1 along the front end of the engine axis, a low-pressure rotating shaft A2 is provided in the middle, and a common low-pressure turbine A3 is provided at the rear end; the common low-pressure turbine A3 is driven by the gas generated by the two gas generator systems in two different working modes, and then drives the low-pressure rotating shaft A2 and the fan A1 to rotate in turn. After the fan A1 rotates, the airflow will be sucked into the engine by the fan A1 and compressed.

[0012] When the crown turbine mode of the low-pressure rotor system A is adopted, the low-pressure rotor system is provided with a fan A1 along the front end of the engine axis, a low-pressure rotating shaft A2 is provided in the middle, and a low-pressure turbine A3 and a crown turbine A4 are provided in two upper and lower independent ducts at the rear end, wherein the moving blades of the crown turbine A4 are located above the blade crowns 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 are a whole; there is a comb tooth sealing structure at the edge of the blade crown of the moving blade of the low-pressure turbine A3, so that the airflow in the duct where the low-pressure turbine is located and the airflow in the duct where the crown turbine is located are independent of each other; when the mixed exhaust turbofan mode is adopted, the airflow of the main combustion chamber B4 blows the low-pressure turbine A3, and then drives the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1; when the air turbine rocket mode is adopted, the airflow of the rocket combustion chamber C1 blows the crown turbine A4, and then drives the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1.

[0013] Furthermore, when the shared low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, 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 fuel and air burn in the main combustion chamber B4 to generate 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 the air flow into the high-pressure compressor B2 and supercharging it; the main combustion chamber B4 adopts an oxygen-rich air-kerosene constant-pressure combustion system to ensure that when the conventional high-pressure core engine system B is working, the downstream dual-mode afterburner combustion chamber E2 has sufficient oxygen for afterburner combustion.

[0014] Furthermore, when the shared low-pressure turbine scheme of the low-pressure rotor system A is adopted, the conventional high-pressure core engine system B is provided with a front mode selection valve B1 that can be actively adjusted to open and close at the inlet of the high-pressure compressor B2, and a rear mode selection valve B6 that is opened and closed synchronously with the front mode selection valve B1 is provided at the outlet of the high-pressure turbine B5. The conventional high-pressure core engine system B can be completely closed by closing the front mode selection valve B1 and the rear mode selection valve B6 at the same time; when the conventional high-pressure core engine system B is working, the rocket gas generator system C is closed, the front mode selection valve B1 and the rear mode selection valve B6 of the conventional high-pressure core engine system B are opened, and the airflow of the high-pressure turbine B5 flows through the rear mode selection valve B6 to drive the shared low-pressure turbine A3 to rotate, and then drive the front low-pressure rotating shaft A2 and the fan A1 to rotate;

[0015] When the Crown turbine scheme of the low-pressure rotor system A is adopted, the conventional high-pressure core engine system B is provided with a front mode selection valve B1 which can actively adjust the opening and closing at the inlet of the duct. The valve B1 can achieve complete opening / closing control of the duct by synchronously actuating with the rear variable area duct adjustment mechanism D5.

[0016] Further, when the common low-pressure turbine scheme of the low-pressure rotor system A is adopted, the rocket gas generator system C includes a rocket combustion chamber C1, a rocket nozzle C2, and an interstage jet support plate C3;

[0017] When the crown turbine scheme of the low-pressure rotor system A is adopted, the rocket gas generator system C includes a rocket combustion chamber C1 and a rocket nozzle C2;

[0018] The rocket combustion chamber C1 adopts a fuel-rich liquid oxygen-kerosene constant pressure combustion system to control the total temperature at the combustion chamber outlet. At the same time, it can ensure that when the rocket gas generator system C is working, the downstream dual-mode afterburner combustion chamber E2 can complete the afterburner combustion without the need for additional kerosene supply; the rocket nozzle C2 is used to expand the pressure and generate a fuel-rich gas flow with a set flow rate.

[0019] When the shared low-pressure turbine scheme of the low-pressure rotor system A is adopted, the propellant burns in the rocket combustion chamber C1 to produce fuel-rich combustion gas, which enters the shared low-pressure turbine A3 through the opened interstage jet support plate C3, expands in the flow channel to drive the shared low-pressure turbine A3 to do work, and drives the low-pressure rotating shaft A2 and the fan A1 to rotate; when the rocket gas generator system C is working, the conventional high-pressure core engine system B is closed, and the air flow sucked by the fan A1 rotates and enters the outer duct D2 of the sub-duct flow system D; the interstage jet support plate C3 includes a jet outlet of the interstage jet support plate, and the jet outlet is also provided with a mechanism that can actively control the opening and closing. When the rocket gas generator system C is not working, the jet outlet of the interstage jet support plate C3 is closed;

[0020] When the crown turbine scheme of the low-pressure rotor system A is adopted, the rocket combustion chamber C1 burns to produce fuel-rich gas, which flows into the crown turbine A4 to expand and do work. The crown turbine A4 drives the low-pressure rotating shaft A2 and the fan A1 to boost the airflow sucked into the engine; when the rocket gas generator system C is working, the conventional high-pressure core engine system B is closed, and the air flow sucked in by the fan A1 all enters the outer duct D2 of the sub-duct flow system D; the air flow flowing through the outer duct D2 and the fuel-rich gas flow at the outlet of the crown turbine A4 are mixed in the mixer E1; when the rocket gas generator system C is working, the front mode selection valve B1 is closed, and the rear variable area duct adjustment mechanism D5 synchronously adjusts to close the inner duct D3, and the conventional high-pressure core engine system B is closed.

[0021] Furthermore, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, the sub-duct flow system D includes three different flow passages, namely, the outer duct D2, the inner duct D3 and the rocket duct D4; the outer duct D2 and the inner duct D3 are formed by the flow diversion at the diverter ring D1 before the inlet of the high-pressure compressor; the rocket duct D4 is the flow passage where the fuel-rich gas flow generated by the rocket combustion chamber is located;

[0022] When the common low-pressure turbine scheme of the low-pressure rotor system A is adopted, the rocket duct D4 is merged into the inner duct D3 through the interstage jet support plate C3, and finally enters the mixer E1 of the re-ignition thrust-increasing system E;

[0023] When the crown turbine mode of the low-pressure rotor system A is adopted, the inner duct D3 and the rocket duct D4 flow independently, and both of them finally flow into the mixer E1 through the rear variable area duct regulating mechanism D5; 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 adjusts 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 adjust and close the inner duct D3 where the low-pressure turbine A3 is located.

[0024] Furthermore, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, the afterburning thrust-increasing system E includes a mixer E1, a dual-mode afterburner E2, and a tail nozzle E3; the gas flow mixed by the mixer E1 enters the dual-mode afterburner E2; in the mixed exhaust turbofan mode, the gas flow at the outlet of the mixer E1 is a mixed flow of oxygen-rich gas and air, and the dual-mode afterburner E2 will inject kerosene for afterburning combustion; in the air turbine rocket mode, the gas flow at the outlet of the mixer E1 is a mixed flow of fuel-rich gas and air, and the dual-mode afterburner E2 does not inject kerosene additionally, but directly ignites the mixed flow for afterburning combustion; after the afterburning combustion of the dual-mode afterburner E2, high-temperature gas is generated, which enters the geometrically adjustable convergent-divergent structure tail nozzle E3 to expand and accelerate, and discharges the high-temperature and high-speed gas flow to generate thrust.

[0025] When the common low-pressure turbine solution of the low-pressure rotor system A is adopted, the air flow entering the mixer E1 is the air flow flowing through the outer duct D2 and the fuel gas flow flowing through the inner duct D3.

[0026] When the crown turbine scheme of the low-pressure rotor system A is adopted, the airflow entering the mixer E1 is the airflow flowing through the outer duct D2, the inner duct D3 and the rocket duct D4; when working in the mixed exhaust turbofan mode, the rocket duct D4 is closed by the rear variable area duct regulating mechanism D5, and the airflow entering the mixer E1 is the air flow flowing through the outer duct D2 and the oxygen-rich fuel gas flow flowing through the inner duct D3; when working in the air turbine rocket mode, the inner duct D3 is closed by the front mode selection valve B1 and the rear variable area duct regulating mechanism D5, and the airflow entering the mixer E1 is the air flow flowing through the outer duct D2 and the fuel-rich fuel gas flow flowing through the rocket duct D4.

[0027] Furthermore, when the shared low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, 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 conventional gas turbine core engine is working, the downstream dual-mode afterburner combustion chamber E2 has sufficient oxygen for combustion; the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant-pressure combustion system. In order to control the total temperature at the rocket combustion chamber outlet, the rocket combustion chamber C1 adopts rich combustion, and at the same time can ensure that when the rocket gas generator system C is working, the downstream dual-mode afterburner combustion chamber E2 can complete combustion without the need for additional kerosene supply.

[0028] Furthermore, when the shared low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, the dual-mode afterburner combustion chamber E2 has two operating modes: when the engine operates in the mixed exhaust turbofan mode, the dual-mode afterburner combustion chamber E2 can actively control the injection amount to choose to perform or not perform afterburner combustion: when the engine operates in the air turbine rocket mode, the dual-mode afterburner combustion chamber E2 no longer injects fuel, and only achieves secondary combustion through ignition and flame stabilization.

[0029] Furthermore, when the shared turbine mode of the low-pressure rotor system A is adopted, the mixed exhaust turbofan mode adopts a geometrically non-adjustable fan design, and the physical rotation speed of the low-pressure rotor system A is controlled in the speed range of Mach number 0.0~4.0 by adjusting the fuel flow of the main combustion chamber B4 in the mixed exhaust turbofan mode and adjusting the throat area of ​​the tail nozzle E3 in the air turbine rocket mode to control the air flow and thrust of the engine; on the fan A1 characteristic diagram, the fan A1 operating point continuously moves to the low converted speed area as the flight Mach number increases, and the converted flow, compression ratio and efficiency of the fan are all reduced.

[0030] Advantages and effects of the present invention

[0031] 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 0.0~2.5. At the same time, the thrust advantage of the air turbine rocket in the speed range of Mach number 2.5~4.0 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 combustion and energy addition of the rocket combustion chamber 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 2.5~4.0, 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.

[0032] 2. The shared low-pressure turbine scheme of the present invention adopts an inter-stage jet support plate structure, so that the engine can share the low-pressure turbine in the mixed exhaust turbofan mode and the air turbine rocket mode, which significantly reduces the design complexity and structural weight of the engine's rotating parts; at the same time, it provides a more significant process control benefit for the bidirectional mode conversion process between the engine's dual working modes, so that the startup 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.

[0033] 3. The Guansheng turbine scheme of the present invention adopts the Guansheng turbine structure, which reduces the difficulty of structural implementation of the 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 the smooth mode conversion of thrust.

[0034] 4. 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 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

[0035] FIG. 1a is a vertical cross-sectional view of an engine of the common low-pressure turbine solution of the present invention;

[0036] FIG1b is a vertical cross-sectional view of an engine of the Guansheng turbine solution of the present invention;

[0037] Figure 1c A partial vertical cross-sectional view 1 of an engine of the common low-pressure turbine solution of the present invention;

[0038] Figure 1d It is a partial vertical cross-sectional view 1 of the engine of the crown turbine scheme of the present invention;

[0039] Figure 1e A partial vertical cross-sectional view 2 of an engine of the common low-pressure turbine solution of the present invention;

[0040] Figure 1f It is a partial vertical cross-sectional view 2 of the engine of the crown turbine scheme of the present invention;

[0041] FIG2a is a schematic diagram of a mixed exhaust turbofan mode of a shared low-pressure turbine solution of the present invention;

[0042] FIG2 b is a schematic diagram of an air turbine rocket mode of the common low-pressure turbine solution of the present invention;

[0043] FIG2c is a schematic diagram of a mixed exhaust turbofan mode of the Guansheng turbine solution of the present invention;

[0044] FIG2d is a schematic diagram of the air turbine rocket mode of the Guansheng turbine solution of the present invention;

[0045] Figure 3 This is a schematic diagram of how the system thrust changes with the flight Mach number;

[0046] Figure 4 This is a schematic diagram showing the variation of system specific impulse with flight Mach number;

[0047] Figure 5 is the variation of the system inlet air flow rate with the flight Mach number;

[0048] Figure 6 The available pressure ratio of the system tail nozzle varies with the flight Mach number;

[0049] Figure 7 The total temperature of the low-pressure turbine inlet of the system changes with the flight Mach number;

[0050] Figure 8 This is a functional block diagram of the dual-mode variable cycle turbine rocket engine of the present invention; DETAILED DESCRIPTION

[0051] Design principle of the present invention

[0052] 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.

[0053] The reason why the combustion state of the rocket combustion chamber is not affected by the 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 the air is that the main combustion chamber B4 adopts an air-kerosene constant pressure combustion system, and its combustion state is related to the air.

[0054] 2. Difficulties in design of the present invention:

[0055] 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 of the two schemes of the present invention (shared low-pressure turbine scheme, crown turbine scheme) adopts the design method of 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.

[0056] 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 by any engine not working in different working speed domains;

[0057] b. The difficulty of inconsistent thermodynamic cycle design requirements encountered in the series structure layout. Aviation gas turbine engines and air turbine rocket engines with series structure layout need to solve the problem of inconsistent requirements for key thermodynamic cycle design parameters of engine bypass ratio and pressure ratio in their respective advantageous operating speed ranges;

[0058] For the mixed exhaust turbofan mode in each of the two schemes, in order to enable the working mode to accelerate to Mach number 2.5 and to maximize the performance of the engine between Mach number 0 and 2.5, the engine needs to make a reasonable compromise design of 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 number 0 and 2.5, and achieve a high specific impulse to reduce fuel consumption during flight acceleration; at the same time, the speed limit of the engine shaft, the temperature limit of the hot end components and the aerodynamic stability limit of the compression components can also be met.

[0059] For the air-turbine rocket mode in each of the two schemes, 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 thermodynamic 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.

[0060] In short, there are significant differences in the boost capacity and bypass ratio requirements of the compressors of the two modes of engines in the two schemes, and this problem needs to be solved when adopting a series structure layout.

[0061] 2. Solution of the present invention

[0062] Aiming at the problem of inconsistent requirements on the boosting capacity and bypass ratio of the compressors of the two modes of engines in the two schemes, the present invention adopts an inner and outer bypass design method that can take into account the requirements of the two modes at the same time.

[0063] 1) For the mixed exhaust turbofan mode of the shared low-pressure turbine scheme as shown in Figure 1a, the inner ducts B1 and B6 are opened synchronously, so that the high-pressure compressor B2 in the inner duct D3 can supercharge 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 suction 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;

[0064] 2) For the air turbine rocket mode with a shared low-pressure turbine solution, such as Figure 1e , the inner ducts B1 and B6 are closed synchronously, and the airflow at D1 is forced to flow into the outer duct D2. By controlling the rich fuel gas flow rate of the rocket duct D4, the total bypass ratio of the engine reaches a relatively large state. 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, it can meet the engine's low total boost ratio requirements.

[0065] 3) For the air turbine rocket model with a shared low-pressure turbine solution, Figure 1c, it is also necessary to solve the problem that the two systems sharing the low-pressure turbine will occupy more radial space. The present invention adopts an interstage jet support plate to solve the problem that the traditional method will occupy more radial space when the two systems share the low-pressure turbine. Because the two working modes need to share the low-pressure turbine, the rocket duct needs to merge into the inner duct before the low-pressure turbine. If a normal annular double-layer flow channel confluence design is adopted, it will occupy more radial space. The interstage jet support plate design makes full use of the internal space of the support plate structure in the engine support structure, and designs the rocket duct and its outlet in the support plate. By increasing the function of the support plate between the high and low pressure turbines, the radial size of the engine can be controlled as much as possible.

[0066] 4) For the mixed exhaust turbofan mode of the Guansheng turbine scheme, the inner duct is used to synchronize the front mode selection valve B1 and the rear variable area duct adjustment mechanism D5 to achieve full 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 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;

[0067] 5) For the air turbine rocket mode of the Guansheng turbine solution, 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.

[0068] 6) For the air turbine rocket mode of the crown turbine scheme, the method of crown turbine system A4 on the low-pressure turbine A3 blade crown is adopted, 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 start-up and shutdown processes of the two core engines will not cause the low-pressure shaft rotor to stop, providing a feasible basis for the mode conversion with smooth thrust.

[0069] The present invention is further explained below in conjunction with the accompanying drawings:

[0070] A dual-mode variable cycle turbo rocket engine is shown in Figures 1a, 1b, Figure 8As shown, its characteristics are: 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, a sub-duct flow system D and a re-ignition thrust-increasing system E, thereby forming two different working modes of the engine, and 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 a mixed exhaust turbofan mode; when the rocket gas generator system C is working, the engine works in an 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;

[0071] The common low-pressure rotor system A includes a common low-pressure turbine solution for reducing the radial size of the engine and the weight of the low-pressure turbine. Figure 1c As shown in the figure, and the Guansheng turbine scheme for decoupling the low-pressure turbine aerodynamic design is as follows Figure 1d As shown;

[0072] When the common low-pressure turbine scheme of the low-pressure rotor system A is adopted, Figure 1c As shown, the conventional high-pressure core engine system B and the rocket gas generator system C share the low-pressure turbine A3 of the low-pressure rotor system A. The shared low-pressure turbine A3 is driven to rotate by the respective combustion gases of the two sets of gas generators in two different working modes, and then drives the low-pressure rotating shaft A2 and the fan A1 in turn, and then the airflow is sucked into the engine by the fan A1 and compressed;

[0073] Supplementary Note 1

[0074] Common low-pressure turbine solution Figure 2a , 2b As shown, when the mixed exhaust turbofan mode is adopted, Figure 2a , 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, Figure 2b , 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.

[0075] When the crown turbine scheme of the low-pressure rotor system A is adopted, Figure 1dAs shown, the gas flow generated by the main combustion chamber B4 of the conventional high-pressure core engine system B drives the low-pressure turbine A3 to rotate, and the rotation of the low-pressure turbine A3 will simultaneously drive the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1 to rotate; the gas flow generated by the rocket combustion chamber C1 of the rocket gas generator system C drives the crown turbine A4 to rotate, and the rotation of the crown turbine A4 will simultaneously drive the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1 to rotate.

[0076] Supplementary Note 2

[0077] Guansheng turbine solution Figure 2c , 2d As shown, when the mixed exhaust turbofan mode is adopted, Figure 2c , 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 2d , 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.

[0078] Further, when the common low-pressure turbine solution of the low-pressure rotor system A is adopted, Figure 1c The low-pressure rotor system A is provided with a fan A1 at the front end along the engine axis, a low-pressure rotating shaft A2 in the middle, and a common low-pressure turbine A3 at the rear end; the common low-pressure turbine A3 is driven by the gas generated by the two gas generator systems in two different working modes, and then drives the low-pressure rotating shaft A2 and the fan A1 to rotate in turn. After the fan A1 rotates, the airflow will be sucked into the engine by the fan A1 and compressed.

[0079] When the crown turbine mode of the low-pressure rotor system A is adopted, Figure 1d The low-pressure rotor system is provided with a fan A1 at the front end of the engine axis, a low-pressure rotating shaft A2 in the middle, and a low-pressure turbine A3 and a crown turbine A4 in two independent upper and lower ducts at the rear end, wherein the moving blades of the crown turbine A4 are located above the blade crowns 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 are a whole; there is a grate sealing structure at the edge of the blade crown of the moving blade of the low-pressure turbine A3, so that the airflows in the duct where the low-pressure turbine is located and the duct where the crown turbine is located are independent of each other; when the mixed exhaust turbofan mode is adopted, the airflow of the main combustion chamber B4 blows the low-pressure turbine A3, and then drives the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1; when the air turbine rocket mode is adopted, the airflow of the rocket combustion chamber C1 blows the crown turbine A4, and then drives the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1.

[0080] Furthermore, if Figure 1c , 1dAs shown, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, 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 fuel and air are burned in the main combustion chamber B4 to generate 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 the air flow into the high-pressure compressor B2 and supercharging it; the main combustion chamber B4 adopts an oxygen-enriched air-kerosene constant-pressure combustion system, mainly to ensure that when the conventional high-pressure core engine system B is working, the downstream dual-mode afterburner combustion chamber E2 has enough oxygen for afterburner combustion.

[0081] Further, when the common low-pressure turbine solution of the low-pressure rotor system A is adopted, as Figure 1c As shown, the conventional high-pressure core engine system B is provided with a front mode selection valve B1 that can be actively adjusted to open and close at the inlet of the high-pressure compressor B2, and a rear mode selection valve B6 that is opened and closed synchronously with the front mode selection valve B1 is provided at the outlet of the high-pressure turbine B5. The conventional high-pressure core engine system B can be completely closed by closing the front mode selection valve B1 and the rear mode selection valve B6 at the same time; when the conventional high-pressure core engine system B is working, the rocket gas generator system C is closed, the front mode selection valve B1 and the rear mode selection valve B6 of the conventional high-pressure core engine system B are opened, and the airflow of the high-pressure turbine B5 flows through the rear mode selection valve B6 and then drives the common low-pressure turbine A3 to rotate, and then drives the front low-pressure rotating shaft A2 and the fan A1 to rotate;

[0082] When the crown turbine scheme of the low-pressure rotor system A is adopted, Figure 1d , 1f As shown, the conventional high-pressure core engine system B is provided with a front mode selection valve B1 which can actively adjust the opening and closing at the inlet of the duct. The valve B1 can realize the complete opening / closing control of the duct by synchronously actuating with the rear variable area duct adjustment mechanism D5.

[0083] Furthermore, if Figure 1c As shown, when the common low-pressure turbine scheme of the low-pressure rotor system A is adopted, the rocket gas generator system C includes a rocket combustion chamber C1, a rocket nozzle C2, and an interstage jet support plate C3;

[0084] When the crown turbine scheme of the low-pressure rotor system A is adopted, Figure 1d As shown, the rocket gas generator system C includes a rocket combustion chamber C1 and a rocket nozzle C2;

[0085] The rocket combustion chamber C1 adopts a fuel-rich liquid oxygen-kerosene constant pressure combustion system to control the total temperature at the combustion chamber outlet. At the same time, it can ensure that when the rocket gas generator system C is working, the downstream dual-mode afterburner combustion chamber E2 can complete the afterburner combustion without the need for additional kerosene supply; the rocket nozzle C2 is used to expand the pressure and generate a fuel-rich gas flow with a set flow rate.

[0086] When the common low-pressure turbine scheme of the low-pressure rotor system A is adopted, Figure 1c The propellant burns in the rocket combustion chamber C1 to produce fuel-rich combustion gas, which enters the common low-pressure turbine A3 through the opened interstage jet support plate C3, expands in the flow channel to drive the common low-pressure turbine A3 to do work, and drives the low-pressure rotating shaft A2 and the fan A1; when the rocket gas generator system C is working, the conventional high-pressure core engine system B is closed, and the air flow sucked by the fan A1 rotates and enters the outer duct D2 of the sub-duct flow system D; the interstage jet support plate C3 includes a jet outlet of the interstage jet support plate, and the jet outlet is also provided with a mechanism that can actively control the opening and closing. When the rocket gas generator system C is not working, the jet outlet of the interstage jet support plate C3 is closed;

[0087] When the crown turbine scheme of the low-pressure rotor system A is adopted, Figure 1d As shown, the rocket combustion chamber C1 burns to produce fuel-rich gas, which flows into the crown turbine A4 to expand and do work. The crown turbine A4 drives the low-pressure rotating shaft A2 and the fan A1 to boost the airflow sucked into the engine. When the rocket gas generator system C is working, the conventional high-pressure core engine system B is closed, and the air flow sucked in by the fan A1 rotates and all enters the outer duct D2 of the sub-duct flow system D. The air flow flowing through the outer duct D2 and the fuel-rich gas flow at the outlet of the crown turbine A4 are mixed in the mixer E1. When the rocket gas generator system C is working, the front mode selection valve B1 is closed, and the rear variable area duct adjustment mechanism D5 is synchronously adjusted to close the inner duct D3, and the conventional high-pressure core engine system B is closed.

[0088] Furthermore, if Figure 1e As shown, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, the sub-duct flow system D includes three different flow passages, namely the outer duct D2, the inner duct D3 and the rocket duct D4; the outer duct D2 and the inner duct D3 are formed by the flow diversion at the diverter ring D1 before the inlet of the high-pressure compressor; the rocket duct D4 is the flow passage where the fuel-rich gas flow generated by the rocket combustion chamber is located;

[0089] When the common low-pressure turbine scheme of the low-pressure rotor system A is adopted, Figure 1e As shown, the rocket duct D4 passes through the interstage jet support plate C3 and merges into the inner duct D3, and finally enters the mixer E1 of the re-ignition boost system E;

[0090] When the crown turbine mode of the low-pressure rotor system A is adopted, Figure 1f , the inner duct D3 and the rocket duct D4 flow independently, and both of them finally flow into the mixer E1 through the rear variable area duct regulating mechanism D5; 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 low-pressure turbine A3 is located.

[0091] Furthermore, if Figure 1e , 1f As shown, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, the afterburning thrust-increasing system E includes a mixer E1, a dual-mode afterburner E2, and a tail nozzle E3; the gas flow mixed by the mixer E1 enters the dual-mode afterburner E2; in the mixed exhaust turbofan mode, the gas flow at the outlet of the mixer E1 is a mixed flow of oxygen-rich gas and air, and the dual-mode afterburner E2 will inject kerosene for afterburning combustion; in the air turbine rocket mode, the gas flow at the outlet of the mixer E1 is a mixed flow of fuel-rich gas and air, and the dual-mode afterburner E2 does not inject kerosene additionally, but directly ignites the mixed flow for afterburning combustion; after the afterburning combustion of the dual-mode afterburner E2, high-temperature gas is generated, which enters the geometrically adjustable convergent-divergent structure tail nozzle E3 to expand and accelerate, and discharges a high-temperature and high-speed gas flow to generate thrust.

[0092] When the common low-pressure turbine solution of the low-pressure rotor system A is adopted, the air flow entering the mixer E1 is the air flow flowing through the outer duct D2 and the fuel gas flow flowing through the inner duct D3.

[0093] When the crown turbine scheme of the low-pressure rotor system A is adopted, Figure 1f As shown, the airflow entering the mixer E1 is the airflow flowing through the outer duct D2, the inner duct D3 and the rocket duct D4; when working in the mixed exhaust turbofan mode, the rocket duct D4 is closed by the rear variable area duct regulating mechanism D5, and the airflow entering the mixer E1 is the air flow flowing through the outer duct D2 and the oxygen-rich gas flow flowing through the inner duct D3; when working in the air turbine rocket mode, the inner duct D3 is closed by the front mode selection valve B1 and the rear variable area duct regulating mechanism D5, and the airflow entering the mixer E1 is the air flow flowing through the outer duct D2 and the oxygen-rich gas flow flowing through the rocket duct D4.

[0094] Furthermore, if Figure 1c , 1dAs shown, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, 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 conventional gas turbine core is working, the downstream dual-mode afterburner combustion chamber E2 has sufficient oxygen for combustion; the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant-pressure combustion system. In order to control the total temperature at the rocket combustion chamber outlet, the rocket combustion chamber C1 adopts rich combustion, and at the same time can ensure that when the rocket gas generator system C is working, the downstream dual-mode afterburner combustion chamber E2 can complete combustion without the need for additional kerosene supply.

[0095] Furthermore, if Figure 1e , 1f As shown, when the common low-pressure turbine scheme of the low-pressure rotor system A or the crown turbine scheme of the low-pressure rotor system A is adopted, the dual-mode afterburner combustion chamber E2 has two working modes: when the engine operates in the mixed exhaust turbofan mode, the dual-mode afterburner combustion chamber E2 can actively control the injection amount to choose to perform or not perform afterburner combustion: when the engine operates in the air turbine rocket mode, the dual-mode afterburner combustion chamber E2 no longer injects fuel, and only realizes secondary combustion through ignition and flame stabilization.

[0096] Furthermore, when the shared turbine mode of the low-pressure rotor system A is adopted, the mixed exhaust turbofan mode adopts a geometrically non-adjustable fan design, and the physical rotation speed of the low-pressure rotor system A is controlled in the speed range of Mach number 0.0~4.0 by adjusting the fuel flow of the main combustion chamber B4 in the mixed exhaust turbofan mode and adjusting the throat area of ​​the tail nozzle E3 in the air turbine rocket mode to control the air flow and thrust of the engine; on the fan A1 characteristic diagram, the fan A1 operating point continuously moves to the low converted speed area as the flight Mach number increases, and the converted flow, compression ratio and efficiency of the fan are all reduced.

[0097] Example 1: Experimental results of two modes

[0098] 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.

[0099] 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.

[0100] 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.

[0101] Embodiment 2

[0102] The present invention is based on the common low-pressure turbine scheme of the low-pressure rotor system A, and designs a method for changing the engine thrust of the common low-pressure turbine with the flight Mach number, which is characterized by:

[0103] Step 1: operate in mixed exhaust turbofan mode within the flight Mach number range of 0.0 to 2.5;

[0104] Step 2: Use the air turbine rocket mode to operate within the flight Mach number range of 2.5 to 4.0;

[0105] 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.

[0106] Furthermore, the engine in step 1 is converted from the air turbine rocket mode to the mixed exhaust turbofan mode, and the specific process is as follows:

[0107] 1) The engine opens the front mode selection valve B1 and the rear mode selection valve B6 synchronously;

[0108] 2) The conventional high-pressure core system B enters the windmill state and the speed increases;

[0109] 3) After the rotation speed of the conventional high-pressure core engine system B reaches the critical value for ignition, the main combustion chamber B4 ignites and the rocket combustion chamber C1 gradually throttles;

[0110] 4) After the speed of the conventional high-pressure core engine system B reaches the critical value for independent operation, the rocket combustion chamber C1 is closed, the interstage jet support plate C3 is closed, and the dual-mode afterburner combustion chamber E2 is shut down. At this point, the engine completes the mode conversion from air turbine rocket mode to mixed exhaust turbofan mode.

[0111] Furthermore, the specific process of the step 2 of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows:

[0112] 1) Synchronously closing the front mode selection valve B1 and the rear mode selection valve B6, while the rocket combustion chamber C1 starts to ignite and gradually produces rich fuel gas;

[0113] 2) The interstage jet support plate C3 is opened;

[0114] 3) The dual-mode afterburner E2 gradually reduces the fuel supply to control the total residual gas coefficient of the engine;

[0115] 4) When the mass flow rate of the rich fuel gas produced by the rocket combustion chamber C1 reaches a certain level, the fuel supply to the rocket combustion chamber C1 and the dual-mode afterburner combustion chamber E2 is cut off at the same time;

[0116] 5) Gradually and synchronously close the front mode selection valve B1 and the rear mode selection valve B6 until the conventional high-pressure core engine system B is completely closed. At this point, the engine completes the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode.

[0117] 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.

[0118] Embodiment 3

[0119] The present invention is based on the crown turbine scheme of the low-pressure rotor system A, and designs a method for changing the engine thrust based on the crown turbine with the flight Mach number, which is characterized by:

[0120] Step 1: operate in mixed exhaust turbofan mode within the flight Mach number range of 0.0 to 2.5;

[0121] Step 2: Use the air turbine rocket mode to operate within the flight Mach number range of 2.5 to 4.0;

[0122] 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.

[0123] The specific process of switching from the air turbine rocket mode to the mixed exhaust turbofan mode is as follows:

[0124] 1) The engine gradually opens the front mode selection valve B1;

[0125] 2) The rear variable area duct adjustment mechanism D5 gradually expands outward;

[0126] 3) Conventional high-pressure core system B enters the windmill state, and the speed gradually increases;

[0127] 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;

[0128] 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;

[0129] 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.

[0130] The specific process of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows:

[0131] 1) First, gradually close the engine mode selection valve B1;

[0132] 2) The rear variable area duct adjustment mechanism D5 gradually retracts to open the rocket duct D4;

[0133] 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) 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;

[0134] 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;

[0135] 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.

[0136] 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.

[0137] 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 dual-mode variable cycle turbo rocket engine, characterized in that: The engine is provided with two 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 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 rocket combustion chamber (C1) in the rocket gas generator system (C) adopts a fuel-rich liquid oxygen-kerosene constant pressure combustion system to control the total temperature at the combustion chamber outlet; the two independent gas generator systems share a low-pressure rotor system (A), a bypass flow system (D), and a re-ignition thrust-increasing system (E), thereby forming two different working modes of the engine; The sub-duct flow system (D) includes three different flow channels, namely an outer duct (D2), an inner duct (D3) and a rocket duct (D4); the outer duct (D2) and the inner duct (D3) are formed by the flow diversion at the diverter ring (D1) before the high-pressure compressor inlet; the rocket duct (D4) is the flow channel where the fuel-rich gas flow generated by the rocket combustion chamber is located; The two different working modes of the engine are: a mixed exhaust turbofan mode and an air turbine rocket mode; when the conventional high-pressure core engine system (B) is working, the rocket gas generator system (C) is shut down, and the engine works in the mixed exhaust turbofan mode; when the rocket gas generator system (C) is working, the conventional high-pressure core engine system (B) is shut down, the fan (A1) rotates and the air flow sucked in all enters the outer duct (D2) of the sub-duct flow system (D), and the engine works in the air turbine rocket mode; The conventional high-pressure core engine system (B) is provided with a front mode selection valve (B1) that can be actively adjusted to open and close at the inlet of the high-pressure compressor (B2), and a rear mode selection valve (B6) that is opened and closed synchronously with the front mode selection valve (B1) is provided at the outlet of the high-pressure turbine (B5), and the conventional high-pressure core engine system (B) can be completely closed by closing the front mode selection valve (B1) and the rear mode selection valve (B6) at the same time; When the flight Mach number is within the range of 0.0 to 2.5, it uses the mixed exhaust turbofan mode, and when the flight Mach number is within the range of 2.5 to 4.0, it uses the air turbine rocket mode; The common low-pressure rotor system (A) includes a common low-pressure turbine solution for reducing the radial size of the engine and the weight of the low-pressure turbine, and a crown turbine solution for decoupling the aerodynamic design of the low-pressure turbine; When the common low-pressure turbine scheme of the low-pressure rotor system (A) is adopted, the conventional high-pressure core engine system (B) and the rocket gas generator system (C) share the low-pressure turbine (A3) of the low-pressure rotor system (A). The common low-pressure turbine (A3) is driven to rotate by the respective gases of the two gas generators in two different working modes, and then drives the low-pressure rotating shaft (A2) and the fan (A1) in turn, and then the airflow is sucked into the engine by the fan (A1) and compressed; The rocket gas generator system (C) comprises a rocket combustion chamber (C1), a rocket nozzle (C2), and an interstage jet support plate (C3); the rocket duct (D4) passes through the interstage jet support plate (C3) and merges into the inner duct (D2), and finally enters the mixer (E1) of the afterburning thrust increasing system (E); When the crown turbine scheme of the low-pressure rotor system (A) is adopted, the gas flow generated by the main combustion chamber (B4) of the conventional high-pressure core engine system (B) blows the low-pressure turbine (A3) to rotate, and the rotation of the low-pressure turbine (A3) will simultaneously drive the crown turbine (A4), the low-pressure rotating shaft (A2) and the fan (A1); the gas flow generated by the rocket combustion chamber (C1) of the rocket gas generator system (C) blows the crown turbine (A4) to rotate, and the rotation of the crown turbine (A4) will simultaneously drive the low-pressure turbine (A3), the low-pressure rotating shaft (A2) and the fan (A1); the rocket gas generator system (C) includes a rocket combustion chamber (C1) and a rocket nozzle (C2); The inner duct (D3) and the rocket duct (D4) flow independently, and both flow into the mixer (E1) through the rear variable area duct adjustment mechanism (D5).

2. A dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When a common low-pressure turbine solution of a low-pressure rotor system (A) is adopted, the low-pressure rotor system (A) is provided with a fan (A1) at the front end along the engine axis, a low-pressure rotating shaft (A2) in the middle, and a common low-pressure turbine (A3) at the rear end; the common low-pressure turbine (A3) is driven by the gas generated by the two gas generator systems in two different working modes, and then drives the low-pressure rotating shaft (A2) and the fan (A1) to rotate in turn; after the fan (A1) rotates, the airflow is sucked into the engine by the fan (A1) and compressed; When the crown turbine mode of the low-pressure rotor system (A) is adopted, the low-pressure rotor system is provided with a fan (A1) along the front end of the engine axis, a low-pressure rotating shaft (A2) in the middle, and a low-pressure turbine (A3) and a crown turbine (A4) in two upper and lower independent ducts at the rear end, wherein the moving blades of the crown turbine (A4) are located above 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) are a whole; there is a grate sealing structure at the edge of the blade crown of the moving blade of the low-pressure turbine (A3), so that the airflows in the duct where the low-pressure turbine is located and the airflows in the duct where the crown turbine is located are independent of each other; When the mixed exhaust turbofan mode is adopted, the airflow of the main combustion chamber (B4) blows the low-pressure turbine (A3), and then drives the crown turbine (A4), the low-pressure rotating shaft (A2) and the fan (A1); when the air turbine rocket mode is adopted, the airflow of the rocket combustion chamber (C1) blows the crown turbine (A4), and then drives the low-pressure turbine (A3), the low-pressure rotating shaft (A2) and the fan (A1).

3. The dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When the common low-pressure turbine scheme of the low-pressure rotor system (A) or the crown turbine scheme of the low-pressure rotor system (A) is adopted, the fuel and air are burned in the main combustion chamber (B4) to generate combustion gas to drive the high-pressure turbine (B5) to rotate. The rotation of the high-pressure turbine (B5) drives the high-pressure rotating shaft (B3) and the high-pressure compressor (B2) to rotate together, and the air flow is sucked into the high-pressure compressor (B2) and pressurized. The main combustion chamber (B4) adopts an oxygen-enriched air-kerosene constant pressure combustion system, mainly to ensure that when the conventional high-pressure core engine system (B) is working, the downstream dual-mode afterburner combustion chamber (E2) has sufficient oxygen for afterburner combustion.

4. A dual-mode variable cycle turbo rocket engine according to claim 3, characterized in that: When the common low-pressure turbine scheme of the low-pressure rotor system (A) is adopted, the conventional high-pressure core engine system (B) is provided with a front mode selection valve (B1) that can be actively adjusted to open and close at the inlet of the high-pressure compressor (B2), and a rear mode selection valve (B6) that is opened and closed synchronously with the front mode selection valve (B1) is provided at the outlet of the high-pressure turbine (B5), and the conventional high-pressure core engine system (B) can be completely closed by closing the front mode selection valve (B1) and the rear mode selection valve (B6) at the same time; the front mode selection valve (B1) and the rear mode selection valve (B6) of the conventional high-pressure core engine system (B) are opened, and the airflow of the high-pressure turbine (B5) flows through the rear mode selection valve (B6) to drive the common low-pressure turbine (A3) to rotate, and then drive the front low-pressure rotating shaft (A2) and the fan (A1) to rotate; When the Guansheng turbine scheme of the low-pressure rotor system (A) is adopted, the conventional high-pressure core engine system (B) is provided with a front mode selection valve (B1) that can actively adjust the opening and closing at the inlet of the duct. The valve (B1) can achieve full opening / closing control of the duct by synchronously actuating with the rear variable area duct adjustment mechanism (D5).

5. The dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When the crown turbine scheme of the low-pressure rotor system (A) is adopted, the rocket gas generator system (C) includes a rocket combustion chamber (C1) and a rocket nozzle (C2); The dual-mode afterburner chamber (E2) downstream of the rocket combustion chamber (C1) can complete afterburner combustion without the need for additional kerosene supply; the rocket nozzle (C2) is used for pressure expansion and generating a fuel-rich gas flow with a set flow rate; When the common low-pressure turbine scheme of the low-pressure rotor system (A) is adopted, the propellant burns in the rocket combustion chamber (C1) to generate fuel-rich combustion gas, which enters the common low-pressure turbine (A3) through the opened interstage jet support plate (C3), expands in the flow channel to drive the common low-pressure turbine (A3) to do work, and drives the low-pressure rotating shaft (A2) and the fan (A1); the interstage jet support plate (C3) includes a jet outlet of the interstage jet support plate, and the jet outlet is also provided with a mechanism that can be actively controlled to open and close. When the rocket gas generator system (C) is not working, the jet outlet of the interstage jet support plate (C3) is closed; When the crown turbine scheme of the low-pressure rotor system (A) is adopted, the rocket combustion chamber (C1) burns to produce fuel-rich gas, which flows into the crown turbine (A4) to expand and do work, and the crown turbine (A4) drives the low-pressure rotating shaft (A2) and the fan (A1) to boost the airflow sucked into the engine; when the rocket gas generator system (C) is working, the conventional high-pressure core engine system (B) is closed, and the air flow sucked in by the fan (A1) all enters the outer duct (D2) of the sub-duct flow system (D); the air flow flowing through the outer duct (D2) and the fuel-rich gas flow at the outlet of the crown turbine (A4) are mixed in the mixer (E1); when the rocket gas generator system (C) is working, the front mode selection valve (B1) is closed, and the rear variable area duct adjustment mechanism (D5) is synchronously adjusted to close the inner duct (D3), and the conventional high-pressure core engine system (B) is closed.

6. The dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When the crown turbine mode of the low-pressure rotor system (A) is adopted; 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) adjusts 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 adjust and close the inner duct (D3) where the low-pressure turbine (A3) is located.

7. The dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When the common low-pressure turbine scheme of the low-pressure rotor system (A) or the crown turbine scheme of the low-pressure rotor system (A) is adopted, the afterburning thrust-increasing system (E) comprises a mixer (E1), a dual-mode afterburner (E2), and a tail nozzle (E3); the gas flow mixed by the mixer (E1) enters the dual-mode afterburner (E2); in the mixed exhaust turbofan mode, the gas flow at the outlet of the mixer (E1) is a mixed flow of oxygen-rich gas and air, and the dual-mode afterburner (E2) injects kerosene for afterburning combustion; in the air turbine rocket mode, the gas flow at the outlet of the mixer (E1) is a mixed flow of fuel-rich gas and air, and the dual-mode afterburner (E2) does not inject kerosene additionally, but directly ignites the mixed flow for afterburning combustion; after the afterburning combustion of the dual-mode afterburner (E2), the high-temperature gas generated enters the geometrically adjustable convergent-divergent tail nozzle (E3) to expand and accelerate, and discharges the high-temperature and high-speed gas flow to generate thrust; When the common low-pressure turbine solution of the low-pressure rotor system (A) is adopted, the air flow entering the mixer (E1) is the air flow flowing through the outer duct (D2) and the gas flow flowing through the inner duct (D3); When the crown turbine scheme of the low-pressure rotor system (A) is adopted, the airflow entering the mixer (E1) is the airflow flowing through the outer duct (D2), the inner duct (D3) and the rocket duct (D4); when the mixed exhaust turbofan mode is used, the rocket duct (D4) is closed by the rear variable area duct regulating mechanism (D5), and the airflow entering the mixer (E1) is the air flow flowing through the outer duct (D2) and the oxygen-rich gas flow flowing through the inner duct (D3); when the air turbine rocket mode is used, the inner duct (D3) is closed by the front mode selection valve (B1) and the rear variable area duct regulating mechanism (D5), and the airflow entering the mixer (E1) is the air flow flowing through the outer duct (D2) and the fuel-rich gas flow flowing through the rocket duct (D4).

8. The dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When the common low-pressure turbine scheme of the low-pressure rotor system (A) or the crown turbine scheme of the low-pressure rotor system (A) is adopted, 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 conventional gas turbine core engine is working, the downstream dual-mode afterburner combustion chamber (E2) has sufficient oxygen for combustion; the rocket combustion chamber (C1) adopts a liquid oxygen-kerosene constant-pressure combustion system. In order to control the total temperature at the rocket combustion chamber outlet, the rocket combustion chamber (C1) adopts rich combustion, and at the same time, it can be ensured that when the rocket gas generator system (C) is working, the downstream dual-mode afterburner combustion chamber (E2) can complete combustion without the need for additional kerosene supply.

9. The dual-mode variable cycle turbo rocket engine according to claim 3, characterized in that: When a common low-pressure turbine scheme of a low-pressure rotor system (A) or a crown turbine scheme of a low-pressure rotor system (A) is adopted, the dual-mode afterburner combustion chamber (E2) has two operating modes: when the engine operates in a mixed exhaust turbofan mode, the dual-mode afterburner combustion chamber (E2) can actively control the amount of fuel injection to choose whether to perform afterburner combustion; when the engine operates in an air turbine rocket mode, the dual-mode afterburner combustion chamber (E2) no longer injects fuel, and only realizes secondary combustion through ignition and flame stabilization.

10. The dual-mode variable cycle turbo rocket engine according to claim 1, characterized in that: When the shared turbine mode of the low-pressure rotor system (A) is adopted, the mixed exhaust turbofan mode adopts a geometrically non-adjustable fan design, and the physical speed of the low-pressure rotor system (A) is controlled in the speed range of Ma 0.0~4.0 by adjusting the fuel flow of the main combustion chamber (B4) in the mixed exhaust turbofan mode and adjusting the throat area of ​​the tail nozzle (E3) in the air turbine rocket mode to control the air flow and thrust of the engine; on the fan (A1) characteristic diagram, the fan (A1) operating point continuously moves to the low converted speed area as the flight Mach number increases, and the converted flow, compression ratio and efficiency of the fan are reduced.

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