Common low-pressure turbine variable cycle turbine rocket engine and thrust realization method thereof

Abstract

The present invention discloses a common low-pressure turbine variable cycle turbo rocket engine and a method for realizing thrust thereof. The engine is provided with two sets of independent gas generator systems: a conventional high-pressure core engine system and a rocket gas generator system. The two sets share a set of low-pressure rotor system, a branch duct flow system, and a re-ignition thrust increasing system, forming two different working modes of the engine: a mixed exhaust turbo fan mode and an air turbo rocket mode. The method comprises: when the engine accelerates from a flight Mach number below 2.5 to a flight Mach number above 2.5, the mixed exhaust turbo fan mode is switched to an air turbo 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 turbo rocket mode is switched to a mixed exhaust turbo fan mode. The present invention adopts a common low-pressure turbine scheme. When the flight Mach number reaches 2.5, the mixed exhaust turbo fan mode is switched to the air turbo rocket mode, and the steady-state thrust of the engine increases sharply.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace vehicle propulsion, and in particular relates to a common low-pressure turbine variable cycle turbine rocket engine and a method for achieving thrust thereof. 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 common low-pressure turbine variable cycle turbine rocket engine and a method for achieving thrust thereof, with the aim of solving the problem that conventional gas turbine engines suffer from insufficient cycle heating under high operating Mach number conditions and are difficult to operate and generate sufficient thrust within the speed range of Mach numbers 2.5 to 4.0.

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

[0007] A variable cycle turbo rocket engine with a shared low-pressure turbine, 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 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 engine receives an aircraft signal: the current flight Mach number is within the range of 0.0 to 2.5, the engine uses the mixed exhaust turbofan mode to work, and when the engine receives an aircraft signal: the current flight Mach number is within the range of 2.5 to 4.0, the engine uses the air turbine rocket mode to work; the two sets of independent gas generator systems, one system is turned on when the other system is turned off;

[0008] Its characteristics are: the low-pressure turbine A3 of the low-pressure rotor system A is a common turbine for two sets of gas generator systems. When working in the air turbine rocket mode, the propellant burns in the rocket combustion chamber C1 of the rocket gas generator system C to produce rich fuel gas, driving the downstream common low-pressure turbine A3 with the rotating fan A1 to inhale and compress the air flow. The rich fuel gas flow and the air flow are mixed and ignited for the second time in the re-ignition thrust-increasing system E, and the gas is discharged at a high speed to generate thrust.

[0009] Furthermore, the low-pressure rotor system A is shared by two independent gas generator systems, which are 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 respective combustion gases in two different working modes, and after being driven, it drives the low-pressure rotating shaft A2 to rotate, and the low-pressure rotating shaft A2 in turn drives the fan A1 to rotate. After the fan A1 rotates, the airflow will be sucked into the entire flow channel by the fan and compressed.

[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; 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 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 high-pressure compressor B2, and a rear mode selection valve B6 that is opened and closed synchronously with the front mode selection valve at the outlet of the high-pressure turbine B5, by simultaneously closing the front mode selection valve B1 and the rear mode selection valve The mode selection valve B6 can completely shut down the conventional high-pressure core engine system B; when the conventional high-pressure core engine system B is working, the rocket gas generator system C is closed, and the rear mode selection valve B6 of the conventional high-pressure core engine system B is opened, and the airflow of the high-pressure turbine B5 flows through the rear mode selection valve B6, and the airflow drawn out from the rear mode selection valve (B6) 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; 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.

[0011] Furthermore, the rocket gas generator system C includes a rocket combustion chamber C1, a rocket nozzle C2, and an interstage jet support plate C3; the rocket nozzle C2 is used to expand the pressure and generate a fuel-rich gas flow with a set flow rate; the propellant burns in the rocket combustion chamber C1 to generate fuel-rich 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 then drives the fan A1 to rotate through the low-pressure rotating shaft A2. When the rocket gas generator system C is working, the conventional high-pressure core engine system B is closed, and the fan A1 rotates to inhale All the air flow 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 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; the rocket combustion chamber C1 adopts a fuel-rich liquid oxygen-kerosene constant pressure combustion system to control the total temperature of the combustion chamber outlet, 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 the afterburner combustion without the need for additional kerosene supply.

[0012] Furthermore, the afterburning thrust-increasing system E includes a mixer E1, a dual-mode afterburner combustion chamber E2, and a tail nozzle E3; the mixer E1 is used to mix the air flow entering from the outer duct D2 and the gas flow entering from the inner duct D3, and the mixed air flow enters the dual-mode afterburner combustion chamber E2 for ignition and combustion, and is finally discharged at high speed through the tail nozzle E3 to generate thrust.

[0013] Furthermore, the sub-duct flow system D includes three different flow channels, 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 diversion at the diverter ring D1 before the high-pressure compressor inlet, and are finally mixed into a single airflow in the mixer E1; the rocket duct D4 is the flow channel where the fuel-rich gas flow generated by the rocket combustion chamber is located, which passes through the interstage jet support plate C3 and merges into the inner duct, and finally enters the mixer E1 of the re-ignition thrust increase system E.

[0014] Furthermore, the dual-mode afterburner E2 has two operating modes: when the engine operates in the mixed exhaust turbofan mode, the dual-mode afterburner E2 can actively control the amount of fuel injection to choose to perform or not perform afterburner combustion; when the engine operates in the air turborocket mode, the dual-mode afterburner E2 no longer injects fuel, and only achieves secondary combustion through ignition and flame stabilization; the tail nozzle E3 is a geometrically adjustable convergent-divergent nozzle.

[0015] Furthermore, in the air turbine rocket mode, by controlling the mass mixing ratio of the oxidant and the reductant of the liquid oxygen-kerosene combustion system in the rocket combustion chamber C1, the total inlet temperature of the common low-pressure turbine A3 is maintained at 1200-1500K to ensure that the total inlet temperature of the common low-pressure turbine A3 in the two working modes is adapted to each other; when operating in the air turbine rocket mode, the mass flow rate of the fuel-rich gas generated by the rocket gas generator system C is the mass flow rate of all propellants.

[0016] Furthermore, the mixed exhaust turbofan mode adopts a geometrically non-adjustable fan design. 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, the physical speed of the fan A1 can be actively adjusted according to performance requirements within the speed range of Ma 0.0 to 4.0; on the fan characteristic diagram, the fan operating point continuously moves to the low reduced speed area as the flight Mach number increases, and the fan's reduced flow, compression ratio and efficiency all decrease; before and after the mode conversion, since the total pressure at the outlet of the low-pressure turbine adapted to the operation of the conventional gas turbine core engine is lower than the total pressure at the outlet of the low-pressure turbine adapted to the operation of the rocket engine, the fan operating point of the air turbine rocket mode is closer to the surge boundary than that of the mixed exhaust turbofan mode, the compression ratio is larger, and the reduced flow is smaller.

[0017] Furthermore, the high-pressure compressor B2 part of the conventional high-pressure core engine system B adopts an axial flow compressor design. When actually determining the specific configuration, a centrifugal compressor design can also be adopted to reduce the axial length of the core engine, while increasing the circumferential space of the turbine flow section to facilitate the arrangement of the rocket gas generator system C.

[0018] A method for varying the thrust of engines sharing a low-pressure turbine with flight Mach number, characterized by:

[0019] Step 1: operate in a mixed exhaust turbofan mode within a flight Mach number range of 0.0 to 2.5;

[0020] Step 2: operate in air turbine rocket mode within the range of flight Mach number 2.5 to 4.0;

[0021] 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.

[0022] 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:

[0023] 1) The engine opens the front mode selection valve B1 and the rear mode selection valve B6 synchronously;

[0024] 2) Conventional high-pressure core system B enters windmill state and the speed increases;

[0025] 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;

[0026] 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.

[0027] Furthermore, the specific process of the step 2 of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows:

[0028] 1) The front mode selection valve B1 and the rear mode selection valve B6 are closed synchronously, and the rocket combustion chamber C1 starts to ignite and gradually produces rich fuel gas;

[0029] 2) The interstage jet support plate C3 is opened;

[0030] 3) The dual-mode afterburner E2 gradually reduces the fuel supply to control the total residual gas coefficient of the engine;

[0031] 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;

[0032] 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.

[0033] 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.

[0034] Advantages and effects of the present invention

[0035] 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.

[0036] 2. The interstage jet support plate structure allows the engine to share the low-pressure turbine in the mixed exhaust turbofan mode and the air turbine rocket mode, significantly reducing 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 two-way mode conversion process between the engine's dual working modes, 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 smooth thrust mode conversion.

[0037] 3. The present invention shares a low-pressure turbine scheme, and 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 to construct a combined cycle engine scheme that meets the requirements of wide-speed range 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

[0038] Figure 1-1 It is a vertical cross-sectional view of the structure of a variable cycle turbo rocket engine with a common low-pressure turbine of the present invention;

[0039] Figure 1-2 One of the partial structural cross-sectional views of the variable cycle turbo rocket engine with a common low-pressure turbine;

[0040] Figure 1-3 The second vertical cross-sectional view of the structure of the variable cycle turbo rocket engine with a common low-pressure turbine of the present invention;

[0041] Figure 2-1 It is a flow diagram of a mixed exhaust turbofan mode of a variable cycle turbo rocket engine sharing a low-pressure turbine of the present invention;

[0042] Figure 2-2 Schematic diagram of the flow of the variable cycle rocket engine with a common low-pressure turbine in the present invention in the air turbine rocket mode

[0043] Figure 3 This is a schematic diagram of how the system thrust changes with the flight Mach number;

[0044] Figure 4 This is a schematic diagram of the system specific impulse changing with the flight Mach number.

[0045] Figure 5 is the variation of the system inlet air flow rate with the flight Mach number;

[0046] Figure 6 The available pressure ratio of the system tail nozzle varies with the flight Mach number;

[0047] Figure 7 The total temperature at the inlet of the system's low-pressure turbine varies with the flight Mach number. DETAILED DESCRIPTION

[0048] Design principle of the present invention

[0049] 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.

[0050] 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.

[0051] 2. Difficulties in design of the present invention:

[0052] 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. 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.

[0053] 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;

[0054] 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;

[0055] 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.

[0056] 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.

[0057] 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.

[0058] 3. Solution of the present invention

[0059] The present invention aims to solve the problem of inconsistent requirements on the boosting capacity and bypass ratio of the compressors of two-mode engines, and adopts an internal and external bypass design method that can take into account the requirements of the two modes at the same time.

[0060] 1) For the mixed exhaust turbofan mode sharing the low-pressure turbine, 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 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;

[0061] 2) For the air turbine rocket mode with a shared low-pressure turbine, 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 requirement.

[0062] 3) For the air turbine rocket mode that shares the low-pressure turbine, 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 of two systems sharing the low-pressure turbine will occupy more radial space. 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.

[0063] Based on the above invention principle, the present invention designs a variable cycle turbine rocket engine with a common low-pressure turbine. 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 engine receives an aircraft signal: the mixed exhaust turbofan mode is used when the current flight Mach number is within the range of 0.0 to 2.5, and when the engine receives an aircraft signal: the air turbine rocket mode is used when the current flight Mach number is within the range of 2.5 to 4.0; the two sets of independent gas generator systems, when one system is turned on, the other system is turned off;

[0064] Its characteristics are: the low-pressure turbine A3 of the low-pressure rotor system A is a common turbine for two sets of gas generator systems. When working in the air turbine rocket mode, the propellant burns in the rocket combustion chamber C1 of the rocket gas generator system C to produce rich fuel gas, driving the downstream common low-pressure turbine A3 with the rotating fan A1 to inhale and compress the air flow. The rich fuel gas flow and the air flow are mixed and ignited for the second time in the re-ignition thrust-increasing system E, and the gas is discharged at a high speed to generate thrust.

[0065] Supplementary Note 1 :

[0066] The present invention has two systems: a conventional high-pressure core engine system B (B system for short) and a rocket gas generator system C (C system for short). The two systems receive signals from an aircraft respectively. The aircraft is provided with a sensor for measuring the Mach number. When the Mach number of the aircraft signal received by the B system is in the range of 0.0 to 2.5, the B system is started; when the Mach number of the aircraft signal received by the C system is in the range of 2.5 to 4.0, the C system is started.

[0067] Supplementary Note 2

[0068] Common low-pressure 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 A is shared by two independent gas generator systems, which are 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 respective combustion gases in two different working modes, and after being driven, it drives the low-pressure rotating shaft A2 to rotate, and the low-pressure rotating shaft A2 in turn drives the fan A1 to rotate, and after the fan A1 rotates, the airflow will be sucked into the entire flow channel by the fan and compressed.

[0070] Supplementary Note 3 :

[0071] like Figure 1-2 As shown, the sharing means that the combustion chamber B4 of system B and the combustion chamber C1 of system C share the low-pressure turbine A3. The airflow led out by the rear mode selection valve (B6) of system B drives the shared low-pressure turbine A3 to rotate through the open interstage jet support plate C3 (B6 is closed at this time).

[0072] Furthermore, if Figure 1-2As 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; 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 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 high-pressure compressor B2, and a rear mode selection valve B6 that is synchronized with the front mode selection valve to open and close at the outlet of the high-pressure turbine B5, by closing the front mode selection valve B1 and the rear mode selection valve B6 at the same time. The rear mode selection valve B6 can completely shut down the conventional high-pressure core engine system B; when the conventional high-pressure core engine system B is working, the rocket gas generator system C is closed, and the rear mode selection valve B6 of the conventional high-pressure core engine system B is opened, and the airflow of the high-pressure turbine B5 flows through the rear mode selection valve B6, and the airflow led out from the rear mode selection valve (B6) 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; 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.

[0073] Supplementary Note 4:

[0074] Only one of the B and C systems can be turned on at the same time while the other is turned off. When the B system is turned off, the front mode selection valve B1 and the rear mode selection valve B6 are closed at the same time. When the C system is turned off, the jet outlet of the interstage jet support plate C3 is closed.

[0075] Furthermore, if Figure 1-2As shown, the rocket gas generator system C includes a rocket combustion chamber C1, a rocket nozzle C2, and an interstage jet support plate C3; the rocket nozzle C2 is used to expand the pressure and generate a fuel-rich gas flow with a set flow rate; the propellant burns in the rocket combustion chamber C1 to generate fuel-rich 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 then drives the fan A1 to rotate through the low-pressure rotating shaft A2. When the rocket gas generator system C is working, the conventional high-pressure core engine system B is closed, and the fan A1 rotates to inhale All the air flows enter 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 capable of controlling 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; the rocket combustion chamber C1 adopts a fuel-rich liquid oxygen-kerosene constant pressure combustion system to control the total temperature of the combustion chamber outlet, 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 afterburner combustion without the need for additional kerosene supply.

[0076] Furthermore, if Figure 1-3 As shown, the afterburner thrust increasing system E includes a mixer E1, a dual-mode afterburner combustion chamber E2, and a tail nozzle E3; the mixer E1 is used to mix the air flow entering from the outer duct D2 and the gas flow entering from the inner duct D3, and the mixed air flow enters the dual-mode afterburner combustion chamber E2 for ignition and combustion, and is finally discharged at high speed through the tail nozzle E3 to generate thrust.

[0077] Furthermore, if Figure 1-3 As shown, the sub-duct flow system D includes three different flow channels, 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 diversion at the diverter ring D1 before the high-pressure compressor inlet, and are finally mixed into a single airflow in the mixer E1; the rocket duct D4 is the flow channel where the fuel-rich gas flow generated by the rocket combustion chamber is located, which passes through the interstage jet support plate C3 and merges into the inner duct, and finally enters the mixer E1 of the re-ignition thrust increase system E.

[0078] Supplementary Note 5:

[0079] like Figure 1-3 As shown in the figure, it is a system with two inlets and two outlets. Although there are three different flow paths, they are merged into the inner channel through the interstage jet support plate C3 and finally converge into two outlets. This design makes the design of the low-pressure rotor system A relatively simple, but the design of the interstage jet support plate C3 is relatively complicated.

[0080] Furthermore, the dual-mode afterburner E2 has two operating modes: when the engine operates in the mixed exhaust turbofan mode, the dual-mode afterburner E2 can actively control the amount of fuel injection to choose to perform or not perform afterburner combustion; when the engine operates in the air turborocket mode, the dual-mode afterburner E2 no longer injects fuel, and only achieves secondary combustion through ignition and flame stabilization; the tail nozzle E3 is a geometrically adjustable convergent-divergent nozzle.

[0081] Furthermore, in the air turbine rocket mode, by controlling the mass mixing ratio of the oxidant and the reductant of the liquid oxygen-kerosene combustion system in the rocket combustion chamber C1, the total inlet temperature of the common low-pressure turbine A3 is maintained at 1200-1500K to ensure that the total inlet temperature of the common low-pressure turbine A3 in the two working modes is adapted to each other; when operating in the air turbine rocket mode, the mass flow rate of the fuel-rich gas generated by the rocket gas generator system C is the mass flow rate of all propellants.

[0082] Furthermore, the mixed exhaust turbofan mode adopts a geometrically non-adjustable fan design. 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, the physical speed of the fan A1 can be actively adjusted according to performance requirements within the speed range of Ma 0.0 to 4.0; on the fan characteristic diagram, the fan operating point continuously moves to the low reduced speed area as the flight Mach number increases, and the fan's reduced flow, compression ratio and efficiency all decrease; before and after the mode conversion, since the total pressure at the outlet of the low-pressure turbine adapted to the operation of the conventional gas turbine core engine is lower than the total pressure at the outlet of the low-pressure turbine adapted to the operation of the rocket engine, the fan operating point of the air turbine rocket mode is closer to the surge boundary than that of the mixed exhaust turbofan mode, the compression ratio is larger, and the reduced flow is smaller.

[0083] Furthermore, the high-pressure compressor B2 part of the conventional high-pressure core engine system B adopts an axial flow compressor design. When actually determining the specific configuration, a centrifugal compressor design can also be adopted to reduce the axial length of the core engine, while increasing the circumferential space of the turbine flow section to facilitate the arrangement of the rocket gas generator system C.

[0084] A method for varying the thrust of engines sharing a low-pressure turbine with flight Mach number, characterized by:

[0085] Step 1: operate in a mixed exhaust turbofan mode within a flight Mach number range of 0.0 to 2.5;

[0086] Step 2: operate in air turbine rocket mode within the range of flight Mach number 2.5 to 4.0;

[0087] 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.

[0088] 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:

[0089] 1) The engine opens the front mode selection valve B1 and the rear mode selection valve B6 synchronously;

[0090] 2) Conventional high-pressure core system B enters windmill state and the speed increases;

[0091] 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;

[0092] 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.

[0093] Furthermore, the specific process of the step 2 of switching from the mixed exhaust turbofan mode to the air turborocket mode is as follows:

[0094] 1) The front mode selection valve B1 and the rear mode selection valve B6 are closed synchronously, and the rocket combustion chamber C1 starts to ignite and gradually produces rich fuel gas;

[0095] 2) The interstage jet support plate C3 is opened;

[0096] 3) The dual-mode afterburner E2 gradually reduces the fuel supply to control the total residual gas coefficient of the engine;

[0097] 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;

[0098] 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.

[0099] 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.

[0100] Example 1: Experimental results of two modes

[0101] 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 4 It 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.

[0102] The present invention is added to the rocket combustion chamber, so that the effect of increasing the inlet flow is as follows 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.

[0103] 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.

[0104] Embodiment 2

[0105] The scheme of the present invention operates in a mixed exhaust turbofan mode within a flight Mach number range of 0.0 to 2.5, and operates in an air turbine rocket mode within a flight Mach number range of 2.5 to 4.0.

[0106] When the scheme of the present invention works in the mixed exhaust turbofan mode, the working principle of the engine is consistent with that of a conventional mixed exhaust turbofan gas turbine engine, such as Figure 2-1 At this time, the front and rear mode selection valves are opened, and the outlet of the flow channel in the interstage jet support plate is closed.

[0107] When the scheme of the present invention works in the air turbine rocket mode, the rocket engine burns to produce oil-rich gas, which enters the low-pressure turbine flow channel through the opened interstage jet support plate, expands in the flow channel to drive the low-pressure turbine to do work, and then drives the fan to rotate, sucking air into the outer duct; the air flow and the oil-rich gas flow are mixed in the mixer, and enter the afterburner to ignite and burn, and finally discharged at high speed through the tail nozzle to generate thrust, such as Figure 2-2 At this time, the front and rear mode selection valves are closed, the flow channel outlet of the interstage jet support plate is opened, and the conventional gas turbine core engine stops rotating.

[0108] When the scheme of the present invention works in the air turbine rocket mode, the mass flow rate of the oil-rich fuel gas generated by the rocket engine is the mass flow rate of the entire propellant.

[0109] In the process of converting from the mixed exhaust turbofan mode to the air turbine rocket mode, the engine first synchronously closes the front mode selection valve and the rear mode selection valve, and at the same time the rocket engine starts to ignite and gradually produces oil-rich gas, and the interstage jet support plate opens; the afterburner chamber gradually reduces the fuel supply to control the total residual gas coefficient of the engine; after the mass flow rate of the oil-rich gas produced by the rocket engine reaches a certain level, the fuel supply to the main combustion chamber and the afterburner chamber is cut off at the same time, and the front mode selection valve and the rear mode selection valve are gradually and synchronously closed until the high-pressure core engine is completely shut down. At this point, the engine completes the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode.

[0110] When the scheme of the present invention switches from the air turbine rocket mode to the mixed exhaust turbofan mode, the engine first gradually and synchronously opens the front mode selection valve and the rear mode selection valve, the high-pressure core engine enters the windmill state, and the speed increases; after the high-pressure compressor speed reaches the ignition critical value, the main combustion chamber ignites, and the rocket engine gradually throttles; after the high-pressure compressor speed reaches the independent working critical value, the rocket engine is shut down, the interstage jet support plate is closed, and the afterburner combustion chamber is extinguished, and the engine completes the mode conversion from the air turbine rocket mode to the mixed exhaust turbofan mode.

[0111] 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 with a common low-pressure turbine, 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 a set of sub-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: a mixed exhaust turbofan mode and an air turborocket 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 turborocket mode; When the engine receives a signal from the aircraft: the current flight Mach number is within the range of 0.0~2.5, it uses the mixed exhaust turbofan mode to work; when the engine receives a signal from the aircraft: the current flight Mach number is within the range of 2.5~4.0, it uses the air turbine rocket mode to work; the two sets of independent gas generator systems, when one system is turned on, the other system is turned off; it is characterized in that: The low-pressure turbine (A3) of the low-pressure rotor system (A) is a common turbine for the two gas generator systems. When the air turbine rocket mode is adopted, the propellant burns in the rocket combustion chamber (C1) of the rocket gas generator system (C) to generate fuel-rich gas, which drives the downstream common low-pressure turbine (A3) with the rotating fan (A1) to inhale and compress the air flow. The fuel-rich gas flow and the air flow are mixed and ignited for the second time in the re-ignition thrust-increasing system (E), and the gas is discharged at a high speed to generate thrust. 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); 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, so that the air flow is sucked into the high-pressure compressor and pressurized; the conventional high-pressure core engine system (B) is also provided with a front mode selection valve (B1) that actively adjusts the opening and closing at the inlet of the high-pressure compressor (B2), and a rear mode selection valve (B6) that opens and closes synchronously with the front mode selection valve at the outlet of the high-pressure turbine (B5), and the conventional high-pressure core engine system (B) is 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 The system (C) is closed, the rear mode selection valve (B6) of the conventional high-pressure core engine system (B) is opened, the airflow of the high-pressure turbine (B5) flows through the rear mode selection valve (B6), and the airflow drawn out from the rear mode selection valve (B6) 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; the main combustion chamber (B4) adopts an oxygen-enriched air-kerosene constant pressure combustion system, which is sufficient 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.

2. According to claim 1, a variable cycle turbo rocket engine with a common low-pressure turbine, characterized in that: The low-pressure rotor system (A) is shared by two independent gas generator systems, and 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 respective gas in two different working modes, and then drives the low-pressure rotating shaft (A2) to rotate, and the low-pressure rotating shaft (A2) in turn drives the fan (A1) to rotate. After the fan (A1) rotates, the airflow will be sucked into the entire flow channel by the fan and compressed.

3. According to claim 1, a variable cycle turbo rocket engine with a common low-pressure turbine, characterized in that: 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 nozzle (C2) is used to expand pressure and generate a fuel-rich gas flow with a set flow rate; the propellant burns in the rocket combustion chamber (C1) to generate a fuel-rich gas, which enters the common low-pressure turbine (A3) through the opened interstage jet support plate (C3), expands in the flow channel, drives the common low-pressure turbine (A3) to do work, and then drives the fan (A1) to rotate through the low-pressure rotating shaft (A2). The rocket gas generator system (C) When working, the conventional high-pressure core engine system (B) is closed, and the air flow sucked by the fan (A1) rotates to enter 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 control 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; the rocket combustion chamber (C1) adopts a fuel-rich liquid oxygen-kerosene constant pressure combustion system to control the total temperature of the combustion chamber outlet, and at the same time ensure that when the rocket gas generator system (C) is working, the downstream dual-mode afterburner combustion chamber (E2) can complete afterburner combustion without the need for additional kerosene supply.

4. According to claim 1, a variable cycle turbo rocket engine with a common low-pressure turbine, characterized in that: The afterburning thrust increasing system (E) comprises a mixer (E1), a dual-mode afterburner combustion chamber (E2), and a tail nozzle (E3); the mixer (E1) is used to mix the air flow entering from the outer duct (D2) and the gas flow entering from the inner duct (D3); the mixed air flow enters the dual-mode afterburner combustion chamber (E2) for ignition and combustion, and is finally discharged at high speed through the tail nozzle (E3) to generate thrust.

5. According to claim 1, a variable cycle turbo rocket engine with a common low-pressure turbine, characterized in that: 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, and are finally mixed into a single airflow in the mixer (E1); the rocket duct (D4) is the flow channel where the fuel-rich gas flow generated by the rocket combustion chamber is located, and it passes through the interstage jet support plate (C3) and merges into the inner duct, and finally enters the mixer (E1) of the re-ignition thrust increase system (E).

6. According to claim 4, a variable cycle turbo rocket engine with a common low-pressure turbine, characterized in that: The dual-mode afterburner (E2) has two operating modes: when the engine operates in the mixed exhaust turbofan mode, the dual-mode afterburner (E2) can actively control the amount of fuel injection to choose whether to perform afterburner combustion; when the engine operates in the air turborocket mode, the dual-mode afterburner (E2) no longer injects fuel, and only achieves secondary combustion through ignition and flame stabilization; the tail nozzle (E3) is a geometrically adjustable convergent-divergent nozzle.

7. The variable cycle turbo rocket engine with a common low-pressure turbine according to claim 1, characterized in that: In the air turbine rocket mode, the total inlet temperature of the common low-pressure turbine (A3) is maintained at 1200~1500K by controlling the mass mixing ratio of the oxidant and the reductant of the liquid oxygen-kerosene combustion system in the rocket combustion chamber (C1), 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; when operating in the air turbine rocket mode, the mass flow rate of the fuel-rich gas generated by the rocket gas generator system (C) is the mass flow rate of all propellants.

8. The variable cycle turbo rocket engine with a common low-pressure turbine according to claim 1, characterized in that: The mixed exhaust turbofan mode adopts a geometrically non-adjustable fan design. 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, the physical speed of the fan (A1) is actively adjusted in the speed range of Ma 0.0~4.0 according to performance requirements; on the fan characteristic diagram, the fan operating point continuously moves to the low reduced speed area as the flight Mach number increases, and the reduced flow, pressure ratio and efficiency of the fan are all reduced; before and after the mode conversion, since the total pressure of the low-pressure turbine outlet adapted to the operation of the conventional gas turbine core engine is lower than the total pressure of the low-pressure turbine outlet adapted to the operation of the rocket engine, the fan operating point of the air turbine rocket mode is closer to the surge boundary than that of the mixed exhaust turbofan mode, the pressure ratio is larger, and the reduced flow is smaller.

9. The variable cycle turbo rocket engine with a common low-pressure turbine according to claim 1, characterized in that: The high-pressure compressor (B2) of the conventional high-pressure core engine system (B) adopts an axial-flow compressor design.

10. A method for varying the engine thrust of a common low-pressure turbine of a variable cycle turbo rocket engine with a common low-pressure turbine according to any one of claims 1 to 9 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: operate in air turbine rocket mode within the flight Mach number range of 2.5 to 4.0; Step 3: switch from mixed exhaust turbofan mode to air turbine rocket mode when the engine accelerates from a flight Mach number below 2.5 to a flight Mach number above 2.5, and switch from air turbine rocket mode to mixed exhaust turbofan mode when the engine decelerates from a flight Mach number above 2.5 to a flight Mach number below 2.

5.

11. The method for changing the thrust of engines sharing a low-pressure turbine with the flight Mach number according to claim 10, characterized in that: The engine in step three is converted from air turbine rocket mode to mixed exhaust turbofan mode, and the specific process is as follows: 1) the engine synchronously opens the front mode selection valve (B1) and the rear mode selection valve (B6); 2) the conventional high-pressure core engine system (B) enters the windmill state and the speed increases; 3) after the speed of the conventional high-pressure core engine system (B) reaches the ignition critical value, the main combustion chamber (B4) ignites and the rocket combustion chamber (C1) gradually throttles; 4) after the speed of the conventional high-pressure core engine system (B) reaches the independent working critical value, the rocket combustion chamber (C1) is closed, the interstage jet support plate (C3) is closed, and the dual-mode afterburner combustion chamber (E2) is turned off. At this point, the engine completes the mode conversion from air turbine rocket mode to mixed exhaust turbofan mode.

12. The method for changing the thrust of engines sharing a low-pressure turbine with the flight Mach number according to claim 10, characterized in that: The engine in step three is converted from the mixed exhaust turbofan mode to the air turbine rocket mode, and the specific process is as follows: 1) The front mode selection valve (B1) and the rear mode selection valve (B6) are closed synchronously, and at the same time, the rocket combustion chamber (C1) starts to ignite and gradually produces rich fuel gas; 2) The interstage jet support plate (C3) is opened; 3) The dual-mode afterburner combustion chamber (E2) gradually reduces the fuel supply to control the total residual gas coefficient of the engine; 4) When the mass flow rate of the rich fuel gas generated 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; 5) The front mode selection valve (B1) and the rear mode selection valve (B6) are gradually and synchronously closed 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.

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