CN115839289A

CN115839289A - Shared low-pressure turbine variable-cycle turbine rocket engine and thrust implementation method thereof

Info

Publication number: CN115839289A
Application number: CN202211262752.6A
Authority: CN
Inventors: 岳连捷; 孟鑫; 王立峰
Original assignee: Institute of Mechanics of CAS
Current assignee: Institute of Mechanics of CAS
Priority date: 2022-10-15
Filing date: 2022-10-15
Publication date: 2023-03-24

Abstract

The invention discloses a shared low-pressure turbine variable-cycle turbine rocket engine and a thrust implementation method thereof, wherein the engine is provided with two sets of mutually independent fuel gas generator systems: the conventional high-pressure core engine system and the rocket gas generator system share a set of low-pressure rotor system, a branch duct flow system and a re-combustion boosting system to form two different working modes of the engine: a hybrid exhaust turbofan mode and an air turbine rocket mode; the method comprises the following steps: the hybrid exhaust turbo fan mode is switched to the air turbo rocket mode when the engine is accelerated from the flying Mach number of 2.5 or less to the flying Mach number of 2.5 or more, and the air turbo rocket mode is switched to the hybrid exhaust turbo fan mode when the engine is decelerated from the flying Mach number of 2.5 or more to the flying Mach number of 2.5 or less. The invention adopts a scheme of sharing the low-pressure turbine, when the flying Mach number reaches 2.5, the mode is switched to the air turbine rocket mode from the mixed exhaust turbine fan mode, and the steady-state thrust of the engine is increased steeply.

Description

Shared low-pressure turbine variable-cycle turbine rocket engine and thrust implementation method thereof

Technical Field

The invention belongs to the field of power propulsion of aerospace aircrafts, and particularly relates to a shared low-pressure turbine variable-cycle turbine rocket engine and a thrust implementation method thereof.

Background

At present, hypersonic aircrafts in the atmosphere have become a hot problem in the field of future aerospace, and a gas turbine engine is considered as an optimal low-speed air suction type power scheme, and the obvious specific impulse advantage of the hypersonic aircrafts can enable the aircrafts using the gas turbine engine as the power scheme to have very high operability and safety in horizontal take-off and landing and low-speed flight.

However, the conventional gas turbine engine is limited by the self-circulation heating limitation and the operating characteristics, and it is difficult to operate in the mach number 2.5-4.0 speed range and generate sufficient thrust, so that the aircraft cannot be effectively flown and accelerated. The reasons for this are three points: firstly, the airflow state of the inlet of the main combustion chamber of the existing gas turbine engine is directly related to the airflow state of the inlet of the engine, namely, the total temperature of the airflow at the inlet of the engine is increased along with the increase of the working Mach number of the engine, so that the total temperature of the airflow at the inlet of the main combustion chamber is correspondingly increased; because the total temperature of 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 working Mach number of the engine is, the smaller the temperature of the main combustion chamber to unit air flow is, the smaller the circulating heating amount of the thermodynamic cycle of the engine is, and the weaker the capability of generating thrust is. The second reason is that the larger the working Mach number of the traditional gas turbine engine is, the larger the total inlet temperature is, and the lower the converted rotating speed of a compression part is; according to the operating characteristics of a general turbomachinery compression part, the lower the reduced rotating speed of the compression part, the poorer the capacity of the compression part to allow air flow, so that the reduced air flow at the inlet of the engine is reduced along with the increase of the operating Mach number; this trend further exacerbates the trend of reduced thrust at high operating mach numbers of the engine. The third reason is that the increase of the working Mach number of the traditional gas turbine engine directly causes the total temperature of the engine to increase in the extension process, and each impeller mechanical part of the engine works under the condition of smaller conversion rotating speed; under the influence of the constraint conditions of insufficient cyclic heating capacity, rotation speed limitation and the like, the engine is difficult to continuously maintain a high physical rotation speed under the condition of high operating Mach number, and the converted rotation speed of each impeller mechanical part of the engine is further reduced; according to the general characteristics of the impeller mechanical parts, the efficiency of the parts is reduced when the converted rotating speed is small; the reduction in efficiency of the various turbomachinery components results in more of the engine cycle heating energy being used to heat the air stream rather than being converted to kinetic energy, which results in increased engine heat rejection losses, reduced thermal and overall efficiency, and further reduced thrust and specific impulse.

In summary, although the theoretical operating mach number upper limit (when the thermal efficiency and the total efficiency are both reduced to 0) of the ideal thermodynamic cycle of the conventional gas turbine engine can reach mach number 4.2, in actual engineering use, due to the complex reasons that each part of the engine cannot reach a completely ideal state, the thrust of the engine is difficult to maintain the flight requirement, or surging occurs on the compression part of the engine, and the like, the operating mach number upper limit of the gas turbine engine in the actual use environment can only reach mach number between 2.5 and 3 generally. This has become a basic consensus in the field of research on aero gas turbine engines.

Disclosure of Invention

The invention provides a shared low-pressure turbine variable-cycle turbine rocket engine and a thrust implementation method thereof, aiming at solving the problems that the traditional gas turbine engine has insufficient cyclic heating capacity under the condition of high operating Mach number, is difficult to operate in the speed range of Mach number 2.5-4.0 and generates enough thrust.

The invention provides the following technical scheme for solving the technical problems:

a variable cycle turbine rocket engine sharing a low pressure turbine is provided with two sets of mutually 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 independent gas generator systems share one low-pressure rotor system A and one bypass flow system D and re-combustion boosting system E, so that two different working modes of the engine are formed, wherein the two different working modes are as follows: a hybrid exhaust turbofan mode and an air turbine rocket mode, the engine operating in the hybrid exhaust turbofan mode when the conventional high pressure core engine system B is operating; when the rocket gas generator system C works, the engine works in an air turbine rocket mode; when the engine receives the airplane signal: the mixed exhaust turbofan mode is used when the current flight Mach number is in the range of 0.0-2.5, and when the engine receives an airplane signal: when the current flight Mach number is within the range of 2.5-4.0, the air turbine rocket mode is used for working; the two sets of gas generator systems are independent from each other, wherein one system is opened while the other system is closed;

the method is characterized in that: when the low-pressure turbine A3 of the low-pressure rotor system A works in an air turbine rocket mode, a propellant burns in a rocket combustion chamber C1 of the rocket gas generator system C to generate rich-fuel gas, the downstream common low-pressure turbine A3 is driven to bring a rotating fan A1 to suck and compress air flow, the rich-fuel gas flow and the air flow are mixed and secondarily ignited in a re-combustion boosting system E, and the gas is discharged at high speed to generate thrust.

Further, the low-pressure rotor system A is shared by two independent gas generator systems, a fan A1 is arranged at the front end of the low-pressure rotor system along the axial direction of the engine, a low-pressure rotating shaft A2 is arranged in the middle of the low-pressure rotor system, and a shared low-pressure turbine A3 is arranged at the rear end of the low-pressure rotor system; the common low-pressure turbine A3 is driven by respective combustion gas in two different working modes, and then drives the low-pressure rotating shaft A2 to rotate, and the low-pressure rotating shaft A2 drives the fan A1 to rotate, and after the fan A1 rotates, the air flow is sucked into the whole flow passage by the fan and compressed.

Further, the conventional high-pressure core 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 oil and air are combusted in the main combustion chamber B4 to generate gas to drive the high-pressure turbine B5 to rotate, the high-pressure turbine B5 rotates to drive the high-pressure rotating shaft B3 and the high-pressure compressor B2 to rotate together, and air flow is sucked into the high-pressure compressor to be pressurized; the conventional high-pressure core machine system B is also provided with a front mode selection valve B1 capable of actively adjusting opening and closing at an inlet of a high-pressure compressor B2, a rear mode selection valve B6 which is synchronously opened and closed with the front mode selection valve is arranged at an outlet of a high-pressure turbine B5, and the conventional high-pressure core machine system B can be completely closed by simultaneously closing the front mode selection valve B1 and the rear mode selection valve B6; when the conventional high-pressure core machine system B works, the rocket gas generator system C is closed, a rear mode selection valve B6 of the conventional high-pressure core machine system B is opened, the airflow of the high-pressure turbine B5 flows through the rear mode selection valve B6, the airflow led out from the rear mode selection valve (B6) drives a common low-pressure turbine A3 to rotate, and then drives a front low-pressure rotating shaft A2 and a fan A1 to rotate; the main combustion chamber B4 adopts an oxygen-enriched air-kerosene constant-pressure combustion system, and mainly ensures that the downstream dual-mode afterburner E2 has enough oxygen for afterburning when the conventional high-pressure core machine system B works.

Further, the rocket gas generator system C comprises a rocket combustion chamber C1, a rocket nozzle C2 and an interstage jet flow support plate C3; the rocket nozzle C2 is used for diffusing and generating a fuel-rich gas flow with set flow rate; the propellant is combusted in a rocket combustion chamber C1 to generate rich-fuel gas, the rich-fuel gas enters a shared low-pressure turbine A3 through an opened interstage jet flow support plate C3, the rich-fuel gas is expanded in a flow channel to drive the shared low-pressure turbine A3 to do work, and then a fan A1 is driven to rotate through a low-pressure rotating shaft A2; the interstage jet flow support plate C3 comprises a jet flow outlet of the interstage jet flow support plate, the jet flow outlet is also provided with a mechanism capable of controlling opening and closing, and when the rocket gas generator system C does not work, the jet flow outlet of the interstage jet flow support plate C3 is closed; the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant pressure combustion system rich in fuel to control the total temperature of the outlet of the combustion chamber, and meanwhile, when the rocket gas generator system C works, the downstream dual-mode afterburner E2 can finish afterburning without additionally supplying kerosene.

Further, the afterburning boosting and thrust boosting system E comprises a mixer E1, a dual-mode afterburner E2 and a tail pipe E3; the mixer E1 is used for mixing 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 E2 for ignition and combustion, and finally the mixed air flow is discharged at a high speed through the tail nozzle E3 to generate thrust.

Further, the sub-duct flow system D includes three different flow passages, which are respectively 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 splitting at a front splitter ring D1 of an inlet of the high-pressure compressor and are finally mixed into single-stranded airflow in a mixer E1; the rocket duct D4 is a flow passage where rich-combustion gas flow generated by the rocket combustion chamber is located, and the rich-combustion gas flow is converged into the inner duct through the interstage jet flow support plate C3 and finally enters the mixer E1 of the afterburning thrust-increasing system E.

Further, there are two modes of operation of the dual mode afterburner E2: when the engine is operating in the mixed-exhaust turbofan mode, the dual-mode afterburner E2 can actively control the amount of fuel injected to select the post-combustion afterburning to be performed or not to be performed: when the engine works in an air turbine rocket mode, the dual-mode afterburner E2 does not spray oil any more, and secondary combustion is realized only through ignition and flame stabilization; the jet pipe E3 is a convergent-divergent nozzle with a geometrically adjustable geometry.

Further, 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 a 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 adaptive; during the air turbine rocket mode of operation, the rocket gas generator system C generates a mass flow of rich fuel gas that is the mass flow of all propellants.

Furthermore, the mixed exhaust turbofan mode adopts a geometric nonadjustable fan design, and the physical rotating speed of the fan A1 can be actively adjusted in the Ma 0.0-4.0 speed range according to performance requirements 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; on a fan characteristic diagram, the working point of the fan continuously moves to a low reduced rotating speed area along with the increase of the flight Mach number, and the reduced flow, the pressure increasing ratio and the efficiency of the fan are all reduced; before and after mode conversion, the total pressure of the low-pressure turbine outlet adapting to the work of the conventional gas turbine core machine is lower than the total pressure of the low-pressure turbine outlet adapting to the work of the rocket engine, so that the working point of the fan in the air turbine rocket mode is closer to a surge boundary, the pressure increase ratio is larger, and the reduced flow is smaller compared with the mixed exhaust turbine fan mode.

Further, the high-pressure compressor B2 part of the conventional high-pressure core machine system B adopts an axial-flow compressor design, and when a specific scheme configuration is actually determined, the centrifugal compressor design can be adopted to reduce the axial length of the core machine, and meanwhile, the circumferential space of a turbine runner section is increased to facilitate the arrangement of the rocket gas generator system C.

A method for changing the thrust of an engine sharing a low-pressure turbine along with the flight Mach number is characterized in that:

step one, working in a mixed exhaust turbofan mode within the range of flight Mach number 0.0-2.5;

step two, working in an air turbine rocket mode within the range of the flight Mach number of 2.5-4.0;

and step three, when the engine is accelerated from the flying Mach number below 2.5 to the flying Mach number above 2.5, the hybrid exhaust turbine fan mode is switched to the air turbine rocket mode, and when the engine is decelerated from the flying Mach number above 2.5 to the flying Mach number below 2.5, the air turbine rocket mode is switched to the hybrid exhaust turbine fan mode.

Further, the engine of the first step is switched from the air turbine rocket mode to the mixed exhaust turbine fan mode, and the specific process is as follows:

1) The engine synchronously opens a front mode selection valve B1 and a rear mode selection valve B6;

2) The conventional high-voltage core machine system B enters a windmill state, and the rotating speed is increased;

3) After the rotating speed of the conventional high-pressure core machine system B reaches an ignition critical value, the main combustion chamber B4 is ignited, and the rocket combustion chamber C1 is gradually throttled;

4) After the rotating speed of the conventional high-pressure core machine system B reaches an independent working critical value, the rocket combustion chamber C1 is closed, the interstage jet flow support plate C3 is closed, the dual-mode afterburner E2 is flamed out, and the engine completes mode conversion from an air turbine rocket mode to a mixed exhaust turbofan mode.

Further, the second step is to switch from the hybrid exhaust turbofan mode to the air turbine rocket mode, and the specific process is as follows:

1) Synchronously closing the front mode selection valve B1 and the rear mode selection valve B6, and simultaneously starting ignition in the rocket combustion chamber C1 and gradually generating rich fuel gas;

2) Opening an interstage jet flow support plate C3;

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

4) When the mass flow of the rich-combustion gas generated by the rocket combustion chamber C1 reaches a certain level, the oil supply of the rocket combustion chamber C1 and the dual-mode afterburner E2 is cut off simultaneously;

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 system B is completely closed, by which time the engine completes the mode transition from the hybrid exhaust turbofan mode to the air turbine rocket mode.

Further, when the engine operates in a mixed exhaust turbofan mode within a range of a flight Mach number of 0.0 to 2.5, thrust and specific impulse characteristics reach set values within a speed range of Ma 0.0 to 2.5; when the engine works in an air turbine rocket mode within the range of the flight Mach number of 2.5-4.0, the steady-state thrust of the engine is increased in a steep rising mode, and the specific impulse of the engine is reduced in a steep falling mode due to the use of the rocket engine.

Advantageous effects of the invention

1. From the overall performance, the scheme of the invention well inherits the good thrust and specific impulse characteristics of the mixed exhaust turbofan engine in the Ma 0.0-2.5 speed domain, and simultaneously makes up the defects that the traditional gas turbine engine has insufficient circulating heating capacity, cannot generate enough thrust and even cannot work in the Mach number 2.5-4.0 speed domain by using the thrust advantage of the air turbine rocket: because the combustion and energy addition of the rocket combustion chamber of the air turbine rocket are not limited by the airflow state at the inlet of the engine any more, the core technical bottlenecks that the thrust is obviously insufficient, the fuel efficiency is obviously reduced and even the normal work can not provide power in the Mach number range of 2.5-4.0 of the conventional gas turbine engine can be effectively solved by converting the engine mode into the air turbine rocket mode, and the air-breathing type variable cycle engine with feasible principle and higher engineering realizability is constructed.

2. An interstage jet flow support plate structure is adopted, so that the engine shares a low-pressure turbine in a mixed exhaust turbine fan mode and an air turbine rocket mode, and the design complexity and the structural weight of a rotating part of the engine are obviously reduced; meanwhile, the method provides more remarkable process control benefits for the bidirectional mode conversion process between the two working modes of the engine, so that the starting and closing processes of the two core machines do not cause the stalling of a low-pressure shaft rotor, and a feasible basis is provided for the mode conversion with stable thrust.

3. The invention shares the scheme of the low-pressure turbine, the thrust in the air turbine rocket mode is increased along with the increase of the flight Mach number, and the power requirement of an aircraft under the high-speed flight condition can be well met; meanwhile, the thrust connection with the air suction type power scheme with higher Mach number is facilitated, and a combined cycle engine scheme meeting the flight use requirement of a wide speed range is constructed. When the flight Mach number reaches 2.5, the scheme of the invention is switched from a mixed exhaust turbofan mode to an air turbine rocket mode, and although the specific impulse of the engine is reduced steeply due to the use of the rocket engine, the steady-state thrust of the engine is increased steeply.

Drawings

FIG. 1-1 is a sectional elevation view of a variable cycle turbine rocket engine configuration sharing a low pressure turbine in accordance with the present invention;

FIGS. 1-2 are sectional elevation views of a portion of the structure of an inventive variable cycle turbine rocket engine sharing a low pressure turbine;

FIGS. 1-3 are sectional elevation views of the second embodiment of the variable cycle turbine rocket engine of the present invention with a common low pressure turbine;

FIG. 2-1 is a schematic flow diagram of a hybrid exhaust turbofan for a variable cycle turbine rocket engine sharing a low pressure turbine according to the present invention;

FIG. 2-2 is a schematic view of the air turbine rocket mode flow of the variable cycle rocket engine sharing a low pressure turbine according to the present invention

FIG. 3 is a schematic representation of system thrust as a function of flight Mach number;

FIG. 4 is a schematic diagram of system specific impulse as a function of flight Mach number.

FIG. 5 is a plot of system inlet air flow rate as a function of flight Mach number;

FIG. 6 is a graph of the pressure ratio available at the nozzle tip of the system as a function of the Mach number of the flight;

FIG. 7 is a graph of total inlet temperature of a low pressure turbine of the system as a function of flight Mach number.

Detailed Description

Design principle of the invention

1. The engine of the invention is added into the rocket combustion chamber for the following reasons:the operating conditions of the main combustion chamber of an original gas turbine engine are directly restricted by the conditions of an engine inlet, and the cyclic heating capacity of the engine is directly limited by the cyclic pressurization ratio and the fuel use. The invention adds an independent rocket combustion chamber on the basis of the original main combustion chamber, and because the rocket combustion chamber does not use air entering the engine as an oxidant, the combustion state of the rocket combustion chamber and the change of the airflow state at the inlet of the circularly heated engine are influenced, namely, the combustion state of the rocket combustion chamber and the working Mach number of the engine do not have direct constraint relation any more. The decoupling of the strong constraint relation solves the core technical bottleneck that the traditional gas turbine engine is insufficient in circulating heating under the condition of high working Mach number, and provides a feasible technical path for realizing the stable work of the engine and generating enough thrust.

The reason why the combustion state of the rocket combustion chamber is not affected by air is that: the total inlet temperature of a common low-pressure turbine A3 is maintained at 1200-1500K by controlling the mass mixing ratio of an oxidant and a reductant of a liquid oxygen-kerosene combustion system in a rocket combustion chamber C1, so as to ensure that the total inlet temperature of the common low-pressure turbine A3 in two working modes is adaptive; it is seen that the combustion state of the combustion chamber is only related to the mass mixing ratio of the oxidant and the reducer of the liquid oxygen-kerosene combustion system and is not related to air. The main combustor combustion conditions and air related causes of the gas turbine combustor are: the main combustion chamber B4 adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-enriched combustion.

2. The design difficulty of the invention is as follows:

because the engine is added with the rocket combustion chamber on the basis of the traditional combustion chamber, two design schemes of parallel structure layout or series structure layout exist in the two combustion chambers and the matched impeller mechanical system. The parallel structure layout is that the main combustion chamber and the impeller machinery matched with the main combustion chamber are respectively divided into two different rotating axes, and the two rotating axes are arranged in parallel. The invention adopts a design method of a mechanical series structure layout of two combustion chambers and matching impellers thereof. The series structure layout is that the main combustion chamber and the matched impeller machinery thereof, and the rocket combustion chamber and the matched impeller machinery thereof are positioned on the same rotating shaft center.

a. The parallel structure layout has the difficulty of 'dead weight'. The aviation gas turbine engine and the air turbine rocket engine which are arranged in a parallel structure bring great difficulty to the combined design of an air inlet and exhaust system used by matching the engines, and the problem of obvious 'dead weight' caused by the fact that any engine in different working speed domains does not work is difficult to solve;

b. the series structure layout meets the difficulty that the design requirements of the thermodynamic cycle are inconsistent. The problems that two engines have different requirements on thermodynamic cycle key design parameters of an engine bypass ratio and a pressure increase ratio in respective advantageous working speed regions need to be solved by adopting an aviation gas turbine engine and an air turbine rocket engine which are arranged in a series structure;

for the mixed exhaust turbine fan mode, in order to accelerate the operation mode to mach number 2.5 and improve the performance of the engine between mach number 0 and mach number 2.5 as much as possible, the engine needs to carry out reasonable compromise design on the thermodynamic cycle design parameters of the engine. The engine adopts a total pressure ratio of about 10-25 and a total bypass ratio of about 0.5-1.5, so that the engine can generate enough thrust between Mach numbers of 0-2.5 and realize higher specific impulse to reduce the fuel consumption in the flight acceleration process; at the same time, the rotational speed limit of the engine's rotating shaft, the temperature limit of the hot end component and the aerodynamic stability limit of the compression component can all be met.

For the air turbine rocket mode, the selection of thermodynamic cycle design parameters for the engine differs significantly from the hybrid exhaust turbofan mode in order to allow the engine to generate sufficient thrust while maintaining a specific impulse that is not too low. By adopting the design of the total bypass ratio of 5-8, on one hand, the total residual oxygen coefficient of the engine can be improved, and the temperature of secondary combustion is increased so as to improve the thrust; on the other hand, the specific impulse of the engine can be effectively improved, and the fuel consumption in the flying acceleration process is reduced as much as possible. The total pressure ratio design of 3-6 is adopted, so that on one hand, the requirement of turbine output power can be reduced, the engine can adopt fewer stages of turbines to drive the fan to do work in a compression manner, and the structural weight of the engine is effectively reduced; on the other hand, the combustion pressure of the rocket combustion chamber can be reduced, and the design difficulty of the rocket combustion chamber and a rocket propellant supply system is reduced.

In a word, the requirements of the supercharging capacity and the bypass ratio of the air compressors of the two-mode engine have great difference, and the problem needs to be solved when the serial structure layout is adopted.

3. Solution of the invention

Aiming at the problem that the requirements of the supercharging capacity and the bypass ratio of the air compressors of the two-mode engine are inconsistent, the invention adopts an inner and outer bypass design method which can simultaneously consider the requirements of the two modes.

1) For a mixed exhaust turbofan mode of a common low-pressure turbine, the inner duct is synchronously opened through B1 and B6, so that the high-pressure compressor B2 in the inner duct D3 can pressurize airflow again, the total pressurization ratio of the engine is higher in the mixed exhaust turbofan mode, meanwhile, the high-pressure compressor B2 has better air suction capacity, more airflow flows into the inner duct D3, the total bypass ratio of the engine is smaller, and the engine better conforms to the trend of the mixed exhaust turbofan mode that the performance of the engine is better;

2) For an air turbine rocket mode sharing a low-pressure turbine, the inner ducts B1 and B6 are synchronously closed, airflow is forced to completely flow into the outer duct D2 at the position D1, the total duct ratio of the engine reaches a larger state by controlling the flow rate of rich-fuel gas of the rocket duct D4, and meanwhile, as the inner duct D3 is closed, the high-pressure compressor B2 stops working, and the supercharging capacity of the fan A1 is lower, the requirement of the low total supercharging ratio of the engine can be met.

3) For the air turbine rocket mode sharing the low-pressure turbine, the problem that the low-pressure turbine shared by the two systems additionally occupies more radial space needs to be solved. The invention adopts the interstage jet flow support plate, and solves the problem that the two systems share the low-pressure turbine to additionally occupy more radial space in the traditional method. Since both modes of operation require a common 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, more radial space is additionally occupied. The interstage jet flow support plate is designed, the internal space of the support plate structure in the engine support structure is fully utilized, and the rocket duct and the outlet of the rocket duct are designed in the support plate. By adding the function of the support plate between the high-pressure turbine and the low-pressure turbine, the radial size of the engine is controlled as much as possible.

Based on the invention principle, the invention designs a variable cycle turbine rocket engine sharing a low-pressure turbine, as shown in figures 1-1, 1-2 and 1-3, the engine is provided with two sets of mutually 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 independent gas generator systems share one low-pressure rotor system A and one bypass flow system D and re-combustion boosting system E, so that two different working modes of the engine are formed, wherein the two different working modes are as follows: a hybrid exhaust turbofan mode and an air turbine rocket mode, the engine operating in the hybrid exhaust turbofan mode when the conventional high pressure core engine system B is operating; when the rocket gas generator system C is in operation, the engine operates in an air turbine rocket mode; when the engine receives the airplane signal: the mixed exhaust turbofan mode is used when the current flight Mach number is in the range of 0.0-2.5, and when the engine receives an airplane signal: when the current flight Mach number is within the range of 2.5-4.0, the air turbine rocket mode is used for working; the two sets of mutually independent gas generator systems, wherein one system is opened and the other system is closed;

the method is characterized in that: when the low-pressure turbine A3 of the low-pressure rotor system A works in an air turbine rocket mode, a propellant burns in a rocket combustion chamber C1 of the rocket gas generator system C to generate rich combustion gas, the downstream common low-pressure turbine A3 is driven to bring a rotating fan A1 to suck and compress air flow, the rich combustion gas flow and the air flow are mixed and secondarily ignited in a re-combustion boosting system E, and the gas is discharged at a high speed to generate thrust.

Supplementary notes 1：

The invention has two sets of systems: the system comprises a conventional high-pressure core machine system B (a system B for short) and a rocket gas generator system C (a system C for short), wherein the two systems respectively receive signals of an airplane, the airplane is provided with a sensor for measuring the Mach number, when the Mach number of the airplane signals received by the system B is in the range of 0.0-2.5, the system B is started, and when the Mach number of the airplane signals received by the system C is in the range of 2.5-4.0, the system C is started.

Supplementary notes 2

The common low pressure turbine scheme is shown in fig. 2-1, 2-2, when a hybrid exhaust turbofan mode is employed, as shown in fig. 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 open; when the air turbine rocket mode is employed, as shown in fig. 2-2, the main combustion chamber B4 of the conventional high pressure core system B is closed and the rocket combustion chamber C1 of the rocket gas generator system C is opened.

Further, as shown in fig. 1-2, 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 in the axial direction of the engine, a low pressure rotating shaft A2 at the middle, and a shared low pressure turbine A3 at the rear end; the common low-pressure turbine A3 is driven by respective combustion gas in two different working modes, and then drives the low-pressure rotating shaft A2 to rotate, and the low-pressure rotating shaft A2 drives the fan A1 to rotate, and after the fan A1 rotates, the air flow is sucked into the whole flow passage by the fan and compressed.

Supplementary notes 3：

As shown in fig. 1-2, the common use is that the combustion chamber B4 of the system B and the combustion chamber C1 of the system C share the low-pressure turbine A3, and the system B drives the common low-pressure turbine A3 to rotate by the airflow led out by the rear mode selection valve (B6) when the rear mode selection valve B6 is opened (at this time, C3 is closed); the C system enters the common low pressure turbine A3 via the open interstage jet support plate C3 (with B6 closed).

Further, as shown in fig. 1-2, the conventional high-pressure core 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 oil and air are combusted in the main combustion chamber B4 to generate gas to drive the high-pressure turbine B5 to rotate, the high-pressure turbine B5 rotates to drive the high-pressure rotating shaft B3 and the high-pressure compressor B2 to rotate together, and air flow is sucked into the high-pressure compressor to be pressurized; the conventional high-pressure core machine system B is also provided with a front mode selection valve B1 capable of actively adjusting opening and closing at an inlet of a high-pressure compressor B2, a rear mode selection valve B6 which is synchronously opened and closed with the front mode selection valve is arranged at an outlet of a high-pressure turbine B5, and the conventional high-pressure core machine system B can be completely closed by simultaneously closing the front mode selection valve B1 and the rear mode selection valve B6; when the conventional high-pressure core machine system B works, the rocket gas generator system C is closed, a rear mode selection valve B6 of the conventional high-pressure core machine system B is opened, the airflow of the high-pressure turbine B5 flows through the rear mode selection valve B6, the airflow led out from the rear mode selection valve (B6) drives a common low-pressure turbine A3 to rotate, and then drives a front low-pressure rotating shaft A2 and a fan A1 to rotate; the main combustion chamber B4 adopts an oxygen-enriched air-kerosene constant-pressure combustion system, and mainly ensures that the downstream dual-mode afterburner E2 has enough oxygen for afterburning when the conventional high-pressure core machine system B works.

Supplementary notes 4:

the system B and the system C can only be one system on and the other system off at the same time. When the system B is closed, the front mode selection valve B1 and the rear mode selection valve B6 are closed simultaneously, and when the system C is closed, the jet outlet of the interstage jet support plate C3 is closed.

Further, as shown in fig. 1-2, the rocket gas generator system C includes a rocket combustion chamber C1, a rocket nozzle C2, an interstage jet plate C3; the rocket nozzle C2 is used for diffusing and generating a fuel-rich gas flow with set flow rate; the propellant is combusted in a rocket combustion chamber C1 to generate rich-fuel gas, the rich-fuel gas enters a shared low-pressure turbine A3 through an opened interstage jet flow support plate C3, the rich-fuel gas is expanded in a flow channel to drive the shared low-pressure turbine A3 to do work, and then a fan A1 is driven to rotate through a low-pressure rotating shaft A2; the interstage jet support plate C3 comprises a jet outlet of the interstage jet support plate, the jet outlet is also provided with a mechanism capable of controlling opening and closing, and the jet outlet of the interstage jet support plate C3 is closed when the rocket gas generator system C does not work; the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant pressure combustion system rich in fuel to control the total temperature of the outlet of the combustion chamber, and meanwhile, when the rocket gas generator system C works, the downstream dual-mode afterburner E2 can finish afterburning without additionally supplying kerosene.

Further, as shown in fig. 1-3, the afterburning boosting system E comprises a mixer E1, a dual-mode afterburner E2, a tail pipe E3; the mixer E1 is used for mixing 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 E2 for ignition and combustion, and finally the mixed air flow is discharged at a high speed through the tail nozzle E3 to generate thrust.

Further, as shown in fig. 1-3, the sub-ducted flow system D includes three different flow passages, respectively 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 splitting at a front splitter ring D1 of an inlet of the high-pressure compressor and are finally mixed into single-stranded airflow in a mixer E1; the rocket duct D4 is a flow passage where rich-combustion gas flow generated by the rocket combustion chamber is located, and the rich-combustion gas flow is converged into the inner duct through the interstage jet flow support plate C3 and finally enters the mixer E1 of the afterburning thrust-increasing system E.

Supplementary notes 5:

1-3, a two inlet and two outlet system, although having three different flow paths, is merged into the inner duct via the interstage jet support plate C3, and finally merged into two outlets, which makes the design of the low pressure rotor system A relatively simple, but the design of the interstage jet support plate C3 is relatively complex

Further, there are two modes of operation of the dual mode afterburner E2: when the engine is operating in the mixed-exhaust turbofan mode, the dual-mode afterburner E2 can actively control the amount of fuel injected to select the post-combustion afterburning to be performed or not to be performed: when the engine works in an air turbine rocket mode, the dual-mode afterburner E2 does not spray oil any more, and secondary combustion is realized only through ignition and flame stabilization; the jet nozzle E3 is a convergent-divergent nozzle with a geometrically adjustable geometry.

Further, 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 a 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 adaptive; during operation in the air turbine rocket mode, the mass flow of the rich fuel gas produced by the rocket gas generator system C is the mass flow of all of the propellant.

Furthermore, the mixed exhaust turbofan mode adopts a geometric nonadjustable fan design, and the physical rotating speed of the fan A1 can be actively adjusted in the Ma 0.0-4.0 speed range according to performance requirements 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; on a fan characteristic diagram, the working point of the fan continuously moves to a low-conversion-speed area along with the increase of the flight Mach number, and the conversion flow, the pressure ratio and the efficiency of the fan are all reduced; before and after mode conversion, the total pressure of the outlet of the low-pressure turbine adapting to the work of the conventional gas turbine core machine is lower than that of the outlet of the low-pressure turbine adapting to the work of the rocket engine, so that the working point of the fan in the air turbine rocket mode is closer to a surge boundary, the pressure increase ratio is higher, and the reduced flow is lower compared with the mixed exhaust turbine fan mode.

Furthermore, the high-pressure compressor B2 part of the conventional high-pressure core machine system B adopts an axial-flow compressor design, and when the specific scheme configuration is actually determined, the centrifugal compressor design can be adopted to reduce the axial length of the core machine, and meanwhile, the circumferential space of a turbine runner section is increased to facilitate the arrangement of the rocket gas generator system C.

secondly, working in an air turbine rocket mode within the range of the flight Mach number of 2.5-4.0;

Further, the engine of the first step is switched from the air turbine rocket mode to the mixed exhaust turbofan mode, and the specific process is as follows:

4) After the rotating speed of the conventional high-pressure core machine system B reaches an independent working critical value, the rocket combustion chamber C1 is closed, the interstage jet flow support plate C3 is closed, and the dual-mode afterburner E2 is flamed out, so that the engine completes mode conversion from an air turbine rocket mode to a mixed exhaust turbofan mode.

2) Opening an interstage jet flow support plate C3;

The first embodiment is as follows: two modes of experimental results

The effect of the invention of adding a rocket combustion chamber to increase the thrust is shown in fig. 3 and 4, fig. 3 shows that the rocket has a large thrust increase when the mach number is more than 2.5, fig. 4 shows that the rocket has a specific impulse decrease when the mach number is more than 2.5, and the comparison of the two figures shows that when the mach number is more than 2.5, the rocket has a large thrust increase although the specific impulse decrease of the rocket combustion chamber is the specific thrust decrease of the fuel or the fuel consumed by the same thrust increases. And when the Mach number is less than 2.5, a mixed exhaust turbofan mode is adopted, the thrust generated by the combustion chamber of the system B is gradually reduced along with the increase of the Mach number and shows a trend of rapid descending when the Mach number is reached, and when the Mach number is more than 2.5, an air turbine rocket mode based on the rocket combustion chamber is adopted, the thrust generated by the system C is gradually increased and is not weakened due to the increase of the Mach number.

The rocket combustion chamber is added, so that the effect of the inlet flow increase is shown in fig. 5, when the Mach number is less than 2,5, because the traditional gas turbine combustion chamber is adopted, the higher the temperature of the main combustion chamber is, the lower the converted rotating speed is, and the lower the converted rotating speed is, the poorer the capacity of allowing the air flow to pass through the combustion chamber is, namely, the lower the inlet flow is. When the Mach number is larger than 2,5, the inlet flow rate of the rocket combustor is not influenced by the air flow and is only related to the combustion material, so when the Mach number is larger than 2,5, the inlet flow rate has an ascending trend. Because the inlet flow is increased, the working efficiency of the engine is improved, the rotating speed of the engine is improved, and the thrust is increased. The effect of the thrust increase is shown in figure 3.

The effect that the total inlet temperature of the low-pressure turbine of the rocket combustor does not rise along with the Mach number is shown in FIG. 7, when the total inlet temperature of the low-pressure turbine of the system B is smaller than the Mach number of 2.5, the total inlet temperature of the low-pressure turbine of the system B rises in direct proportion to the Mach number, and when the total inlet temperature of the low-pressure turbine of the system C is larger than the Mach number of 2.5, the total inlet temperature of the low-pressure turbine of the system C is constant and not in direct proportion to the Mach number.

Example two

The scheme of the invention uses a mixed exhaust turbofan mode to work in the range of the flight Mach number of 0.0-2.5, and uses an air turbine rocket mode to work in the range of the flight Mach number of 2.5-4.0.

The operating principle of the engine when operating in the mixed exhaust turbofan mode is consistent with that of a conventional mixed exhaust turbofan gas turbine engine, as shown in fig. 2-1. At the moment, the front mode selection valve and the rear mode selection valve are opened, and the outlet of the flow channel in the interstage jet flow support plate is closed.

When the scheme of the invention works in an air turbine rocket mode, a rocket engine combusts to generate rich fuel gas, the rich fuel gas enters a low-pressure turbine flow passage through an opened interstage jet flow support plate, the low-pressure turbine is driven to do work by expansion in the flow passage, a fan is driven to rotate, and the sucked air enters an outer bypass; the air flow and the fuel-rich gas flow are mixed in the mixer, enter the afterburner to be ignited and combusted, and are finally discharged at a high speed through the tail nozzle to generate thrust, as shown in the figure 2-2. At the moment, the front mode selection valve and the rear mode selection valve are closed, the outlet of the flow channel in the interstage jet flow support plate is opened, and the conventional gas turbine core engine stops rotating.

When the scheme of the invention works in an air turbine rocket mode, the mass flow of the fuel-rich gas generated by the rocket engine is the mass flow of all propellants.

In the process of switching from a mixed exhaust turbofan mode to an air turbine rocket mode, an engine is firstly synchronously turned off to form a front mode selection valve and a rear mode selection valve, meanwhile, the rocket engine starts ignition and gradually generates rich fuel gas, and an interstage jet flow support plate is opened; the afterburner gradually reduces oil supply to control the total residual gas coefficient of the engine; and after the mass flow of the rich fuel gas generated by the rocket engine reaches a certain level, simultaneously cutting off the oil supply of the main combustion chamber and the afterburner chamber, gradually and synchronously closing the front mode selection valve and the rear mode selection valve until the high-pressure core engine is completely closed, and completing the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode by the engine.

When the air turbine rocket mode is converted into the mixed exhaust turbofan mode, the engine gradually and synchronously opens the front mode selection valve and the rear mode selection valve, the high-pressure core engine enters a windmill state, and the rotating speed is increased; after the rotating speed of the high-pressure compressor reaches an ignition critical value, the main combustion chamber is ignited, and the rocket engine is gradually throttled; after the rotating speed of the high-pressure compressor reaches an independent working critical value, the rocket engine is closed, the interstage jet flow support plate is closed, the afterburner is flamed out, and the mode conversion from the air turbine rocket mode to the mixed exhaust turbofan mode is completed by the engine.

It should be emphasized that the above-described embodiments are merely illustrative of the present invention and are not limiting, since modifications and variations of the above-described embodiments, which are not inventive, may occur to those skilled in the art upon reading the specification, are possible within the scope of the appended claims.

Claims

1. A kind of variable cycle turbine rocket engine sharing the low-pressure turbine, the engine has two sets of independent gas generator systems, one is the conventional high-pressure core machine system (B), another is the rocket gas generator system (C); the two independent gas generator systems share one low-pressure rotor system (A), one branch bypass flow system (D) and the afterburning boosting system (E), so that two different working modes of the engine are formed, wherein the two different working modes are as follows: a hybrid exhaust turbofan mode and an air turbine rocket mode, the engine operating in the hybrid exhaust turbofan mode when the conventional high pressure core system (B) is operating; during operation of the rocket gas generator system (C), the engine operates in an air turbine rocket mode; when the engine receives the signal of the airplane: when the current flight mach number is in the range of 0.0 to 2.5, the mixed exhaust turbofan mode is used for working, and when the engine receives signals of the airplane: the current flight Mach number is 2.5-4.0, and the air turbine rocket mode is used for working; the two sets of mutually independent gas generator systems, wherein one system is opened and the other system is closed;

the method is characterized in that: when the low-pressure turbine (A3) of the low-pressure rotor system (A) works in an air turbine rocket mode, the propellant burns in a rocket combustion chamber (C1) of the rocket gas generator system (C) to generate rich combustion gas, the downstream common low-pressure turbine (A3) with a rotating fan (A1) is driven to suck and compress air flow, the rich combustion gas flow and the air flow are mixed and secondarily ignited in the re-combustion boosting system (E), and the gas is discharged at high speed to generate thrust.

2. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the low-pressure rotor system (A) is shared by two independent gas generator systems, a fan (A1) is arranged at the front end along the axial direction of the engine, a low-pressure rotating shaft (A2) is arranged in the middle of the low-pressure rotor system, and a shared low-pressure turbine (A3) is arranged at the rear end of the low-pressure rotor system; the shared low-pressure turbine (A3) is driven by respective gas in two different working modes, the low-pressure rotating shaft (A2) is driven to rotate after the common low-pressure turbine is driven, the fan (A1) is driven to rotate by the low-pressure rotating shaft (A2), and after the fan (A1) rotates, airflow can be sucked into the whole flow passage by the fan and compressed.

3. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the conventional high-pressure core machine system (B) comprises a high-pressure air compressor (B2), a high-pressure rotating shaft (B3), a main combustion chamber (B4) and a high-pressure turbine (B5); fuel oil and air are combusted in the main combustion chamber (B4) to generate gas to drive the high-pressure turbine (B5) to rotate, the high-pressure turbine (B5) rotates to drive the high-pressure rotating shaft (B3) and the high-pressure compressor (B2) to rotate together, and air flow is sucked into the high-pressure compressor to be pressurized; the conventional high-pressure core machine system (B) is also provided with a front mode selection valve (B1) capable of actively adjusting opening and closing at an inlet of a high-pressure compressor (B2), a rear mode selection valve (B6) which is synchronously opened and closed with the front mode selection valve is arranged at an outlet of a high-pressure turbine (B5), and the conventional high-pressure core machine system (B) can be completely closed by simultaneously closing the front mode selection valve (B1) and the rear mode selection valve (B6); when the conventional high-pressure core engine system (B) works, the rocket gas generator system (C) is closed, a 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 led out from the rear mode selection valve (B6) drives a common low-pressure turbine (A3) to rotate and then drives a front low-pressure rotating shaft (A2) and a fan (A1) to rotate; the main combustion chamber (B4) adopts an oxygen-enriched air-kerosene constant-pressure combustion system, and mainly ensures that the downstream dual-mode afterburner (E2) has enough oxygen for afterburning when the conventional high-pressure core machine system (B) works.

4. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the rocket gas generator system (C) comprises a rocket combustion chamber (C1), a rocket nozzle (C2) and an interstage jet flow support plate (C3); the rocket nozzle (C2) is used for diffusing and generating a fuel-rich gas flow with a set flow rate; the propellant is combusted in a rocket combustion chamber (C1) to generate rich-fuel gas, the rich-fuel gas enters a common low-pressure turbine (A3) through an opened interstage jet flow support plate (C3), the common low-pressure turbine (A3) is driven to do work by expansion in a flow channel, and then a fan (A1) is driven to rotate through a low-pressure rotating shaft (A2), when the rocket gas generator system (C) works, a conventional high-pressure core machine system (B) is closed, and air flow sucked by rotation of the fan (A1) completely enters an outer duct (D2) of a sub-duct flow system (D); the interstage jet flow support plate (C3) comprises a jet flow outlet of the interstage jet flow support plate, the jet flow outlet is also provided with a mechanism capable of controlling opening and closing, and when the rocket gas generator system (C) does not work, the jet flow outlet of the interstage jet flow support plate (C3) is closed; the rocket combustion chamber (C1) adopts a liquid oxygen-kerosene constant pressure combustion system rich in fuel to control the total temperature of the outlet of the combustion chamber, and meanwhile, when the rocket fuel gas generator system (C) works, the downstream dual-mode afterburner (E2) can finish afterburning without additionally supplying kerosene.

5. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the afterburning boosting system (E) comprises a mixer (E1), a dual-mode afterburner (E2) and a tail nozzle (E3); the mixer (E1) is used for mixing 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 (E2) for ignition and combustion, and finally the mixed air flow is discharged at a high speed through the tail nozzle (E3) to generate thrust.

6. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the sub-duct flow system (D) comprises three different flow passages, 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 branching at a splitter ring (D1) in front of an inlet of the high-pressure compressor and are finally mixed into a single-strand airflow in a mixer (E1); the rocket duct (D4) is a flow channel where rich-combustion gas flow generated by the rocket combustion chamber is located, and the rich-combustion gas flow is converged into the inner duct through the interstage jet flow support plate (C3) and finally enters the mixer (E1) of the afterburning and thrust-increasing system (E).

7. The variable cycle turbo-rocket engine sharing a low-pressure turbine of claim 5 wherein: there are two modes of operation of the dual mode afterburner (E2): when the engine is operating in the mixed-exhaust turbofan mode, the dual-mode afterburner (E2) can actively control the amount of fuel injected to select the afterburning to be performed or not to be performed: when the engine works in an air turbine rocket mode, the dual-mode afterburner (E2) does not spray oil any more, and secondary combustion is realized only through ignition and flame stabilization; the jet nozzle (E3) is a convergent-divergent nozzle with adjustable geometry.

8. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: in the air turbine rocket mode, the total inlet temperature of the shared low-pressure turbine (A3) is maintained at 1200-1500K by controlling the mass mixing ratio of an oxidant and a reductant of a liquid oxygen-kerosene combustion system in a rocket combustion chamber (C1), so that the total inlet temperature of the shared low-pressure turbine (A3) in the two working modes is ensured to be adaptive; in the air turbine rocket mode of operation, the rocket gas generator system (C) generates a mass flow of rich fuel gas as the mass flow of all propellants.

9. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the mixed exhaust turbofan mode adopts a geometrically nonadjustable fan design, and the physical rotating speed of the fan (A1) can be actively adjusted within the Ma 0.0-4.0 speed range according to performance requirements by adjusting the fuel flow of a main combustion chamber (B4) in the mixed exhaust turbofan mode and adjusting the throat area of a tail nozzle (E3) in the air turbine rocket mode; on a fan characteristic diagram, the working point of the fan continuously moves to a low-conversion-speed area along with the increase of the flight Mach number, and the conversion flow, the pressure ratio and the efficiency of the fan are all reduced; before and after mode conversion, the total pressure of the low-pressure turbine outlet adapting to the work of the conventional gas turbine core machine is lower than the total pressure of the low-pressure turbine outlet adapting to the work of the rocket engine, so that the working point of the fan in the air turbine rocket mode is closer to a surge boundary, the pressure increase ratio is larger, and the reduced flow is smaller compared with the mixed exhaust turbine fan mode.

10. A variable cycle turbo-rocket engine sharing a low-pressure turbine according to claim 1, wherein: the high-pressure compressor (B2) part of the conventional high-pressure core machine system (B) adopts an axial-flow compressor design, and when the specific scheme configuration is actually determined, the centrifugal compressor design can be adopted to reduce the axial length of the core machine, and meanwhile, the circumferential space of a turbine runner section is increased to facilitate the arrangement of the rocket gas generator system (C).

11. A method of changing the engine thrust with the flight mach number of a common low pressure turbine of a variable cycle turbine rocket engine based on any one of claims 1 to 10, wherein:

12. A method of sharing low pressure turbine engine thrust with flight mach number as recited in claim 11 wherein: the engine of the first step is switched from an air turbine rocket mode to a mixed exhaust turbine fan mode, and the specific process is as follows:

1) The engine synchronously opens a front mode selection valve (B1) and a rear mode selection valve (B6);

2) The conventional high-voltage core machine system (B) enters a windmill state, and the rotating speed is increased;

3) After the rotating speed of the conventional high-pressure core machine system (B) reaches an ignition critical value, the main combustion chamber (B4) is ignited, and the rocket combustion chamber (C1) is gradually throttled;

4) After the rotating speed of the conventional high-pressure core machine system (B) reaches an independent working critical value, the rocket combustion chamber (C1) is closed, the interstage jet flow support plate (C3) is closed, the dual-mode afterburner (E2) is flamed out, and the mode conversion from the air turbine rocket mode to the mixed exhaust turbofan mode is completed by the engine.

13. A method of sharing low pressure turbine engine thrust with flight mach number as recited in claim 11 wherein: and step two, switching from a mixed exhaust turbofan mode to an air turbine rocket mode, wherein the specific process is as follows:

1) Synchronously closing the front mode selection valve (B1) and the rear mode selection valve (B6), and simultaneously starting ignition of the rocket combustion chamber (C1) and gradually generating rich fuel gas;

2) Opening an interstage jet flow support plate (C3);

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

4) When the mass flow of the rich fuel gas generated by the rocket combustion chamber (C1) reaches a certain level, the oil supply of the rocket combustion chamber (C1) and the dual-mode afterburner (E2) is cut off simultaneously;

5) And gradually closing the front mode selection valve (B1) and the rear mode selection valve (B6) synchronously until the conventional high-pressure core machine system (B) is completely closed, so that the engine completes the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode.

14. A method of varying engine thrust with mach number of flight sharing a low pressure turbine in accordance with claim 11, wherein: when the engine works in a mixed exhaust turbofan mode in the range of the flight Mach number of 0.0-2.5, the thrust and specific impulse characteristics reach set values in the speed range of Ma 0.0-2.5; when the engine works in an air turbine rocket mode within the range of the flight Mach number of 2.5-4.0, the steady-state thrust of the engine is increased in a steep rising mode, and the specific impulse of the engine is reduced in a steep falling mode due to the use of the rocket engine.