CN115680940A

CN115680940A - Crown turbine variable-cycle turbine rocket engine and engine thrust implementation method

Info

Publication number: CN115680940A
Application number: CN202211262753.0A
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: 2023-02-03

Abstract

The invention discloses a crown turbine variable cycle turbine rocket engine and an engine thrust implementation method, wherein the engine is provided with two independent fuel gas generator systems: the conventional high-pressure core engine system and the rocket gas generator system share the low-pressure rotor system to form two different working modes: hybrid exhaust turbofan and air turbine rocket modes; the low-pressure rotor system is provided with an integrated low-pressure rotor turbine which is respectively arranged in an inner duct and an outer duct; the method comprises the following steps: the hybrid exhaust turbo fan mode is switched to the air turbo rocket mode when the engine is decelerated from 2.5 or more to 2.5 or less in mach number, and the hybrid exhaust turbo fan mode is switched to the hybrid exhaust turbo fan mode when the engine is decelerated from 2.5 or more in mach number to 2.5 or less in mach number. When the flight Mach number reaches 2.5, the mode of the mixed exhaust turbine fan is switched to the air turbine rocket mode, and the steady-state thrust of the engine is increased steeply.

Description

Crown turbine variable-cycle turbine rocket engine and engine thrust implementation method

Technical Field

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

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 main combustion chamber inlet of the existing gas turbine engine is directly related to the airflow state of the engine inlet, 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 engine thrust reduction at high operating mach numbers. For the third reason, the increase of the working Mach number of the traditional gas turbine engine directly causes the total temperature of the engine to rise in the process of extending, and each impeller mechanical part of the engine works under the condition of smaller converted 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 fundamental consensus in the field of research for aviation gas turbine engines.

Disclosure of Invention

The invention provides a crown turbine variable-cycle turbine rocket engine and an engine thrust implementation method aiming at the defects of the prior art and aims to solve the problems that the traditional gas turbine engine is insufficient in cyclic heating 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 using a coronal turbine system 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; when the conventional high-pressure core system B works, the engine works in a mixed exhaust turbofan mode; when the rocket gas generator system C is in operation, the engine operates in an air turbine rocket mode; when the flying Mach number is within the range of 0.0-2.5, the hybrid exhaust turbofan mode is used for working, and when the flying 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: the low-pressure rotor system A is provided with a low-pressure rotor turbine which is integrally manufactured but is divided into an inner duct and an outer duct, wherein the low-pressure rotor turbine is positioned in the outer rocket duct D4 and is a crown turbine A4, and the low-pressure rotor turbine is positioned in the inner duct D3 and is a low-pressure turbine A3; the shrouded turbine A4 movable blade is fixedly connected above the shroud of the low-pressure turbine A3 movable blade; when the gas drives the low-pressure turbine A3 or the coronally generated turbine A4, the low-pressure rotating shaft A2 and the fan A1 can be simultaneously rotated; when the engine works in the mixed exhaust turbofan mode, the oxygen-enriched gas at the outlet of the conventional core system B drives a low-pressure turbine A3 to rotate and bring a fan A1 to suck and compress air flow; when the engine works in the air turbine rocket mode, the rich combustion gas at the outlet of the rocket gas generator system C drives a crown turbine A4 to suck and compress air flow with a rotating fan A1; the gas flow and the air flow are mixed and secondarily ignited in the afterburning boosting system E, and are discharged at a high speed to generate thrust.

Further, the low-pressure rotor system is composed of a fan A1, a low-pressure turbine A3, a crown turbine A4 and a low-pressure rotating shaft A2; the shrouded turbine A4 movable blade is fixedly connected above a low-pressure turbine A3 movable blade shroud, and the low-pressure turbine A3 movable blade and the shrouded turbine A4 movable blade rotate at the same constant angular speed as a whole; a labyrinth sealing structure is arranged at the edge of a blade crown of a movable blade of the low-pressure turbine A3, so that the airflow of an inner duct D3 where the low-pressure turbine is located and the airflow of a rocket duct D4 where the shrouded turbine is located are independent of each other; when the mixed exhaust turbine fan mode is adopted for work, oxygen-enriched gas flow generated by combustion of the main combustion chamber B4 blows the low-pressure turbine A3 through the high-pressure turbine B5, and the low-pressure turbine A3 rotates and simultaneously drives the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1 to rotate; when the air turbine rocket mode is adopted, the rich combustion gas flow generated by the combustion of the rocket combustion chamber C1 blows the crown turbine A4, and the crown turbine A4 rotates and simultaneously rotates the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1.

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; the main combustion chamber B4 combusts to generate oxygen-enriched 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 partial air flow flowing through the fan A1 is sucked and 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 inner duct inlet, and the front mode selection valve B1 and a rear variable area duct adjusting mechanism D5 synchronously act to realize the control of completely opening/closing the inner duct D3 where the conventional high-pressure core machine system B and the low-pressure turbine A3 are located.

Further, the rocket gas generator system C comprises a rocket combustion chamber C1 and a rocket nozzle C2; the rocket combustion chamber C1 combusts to generate rich fuel gas, the rich fuel gas flows into the crown turbine A4 to expand and do work, and the crown turbine A4 drives the low-pressure shaft to rotate A2, so that the fan A1 pressurizes air flow sucked into the engine.

Further, the bypass-divided flow system D comprises a high-pressure compressor inlet front shunting ring D1, an outer bypass D2, an inner bypass D3, a rocket bypass D4 and a rear variable-area bypass adjusting mechanism D5; the outer duct D2 and the inner duct D3 are formed by splitting at the front splitter ring D1 of the inlet of the high-pressure compressor; the air flow passing through the outer duct D2 finally flows directly into the mixer E1; the airflow passing through the inner duct D3 finally flows into the mixer E1 through the rear variable-area duct adjusting mechanism D5; the rocket duct D4 is a flow passage where rich-combustion gas flow generated by the rocket gas generator system C is located, and the rich-combustion gas flow flows into the mixer E1 through the rear variable-area duct adjusting mechanism D5 after flowing through the crown turbine A4.

Further, the rear variable-area duct adjusting mechanism D5 can control the airflows of the inner duct D3 and the rocket duct D4 at the same time; when the mixed exhaust turbine fan mode is used for working, the rear variable-area duct adjusting mechanism D5 adjusts and closes the rocket duct D4 where the coronal turbine A4 is located; when the air turbine rocket mode is used, the rear variable-area duct adjusting mechanism D5 is matched with the front mode selection valve B1 to adjust and close the inner duct D3 where the conventional high-pressure core machine system B and the low-pressure turbine A3 are located.

Further, the main combustion chamber B4 adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-enriched combustion, so that when the mixed exhaust turbine fan works in a mode, the downstream dual-mode afterburner E2 has enough oxygen to perform afterburning.

Further, the rocket combustion chamber C1 adopts a liquid oxygen-kerosene constant pressure combustion system, in order to control the total temperature of the outlet of the rocket combustion chamber C1, the rocket combustion chamber C1 adopts rich combustion, and meanwhile, when the air turbine rocket mode is used for working, the downstream dual-mode afterburner E2 can finish secondary combustion 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; when operating in the mixed exhaust turbofan mode, the mixer E1 is used for mixing the air flow flowing in the outer duct D2 and the oxygen-enriched gas flow flowing in the inner duct D3; when operating in the air turbine rocket mode, the mixer E1 is used for mixing the air flow flowing in from the outer duct D2 and the rich gas flow flowing in from the rocket duct D4; the airflow mixed by the mixer E1 enters a dual-mode afterburner E2 for afterburning combustion; when the mixed exhaust turbine fan mode works, the dual-mode afterburner E2 can freely adjust to carry out afterburning or not by controlling the supply amount of kerosene; when the air turbine rocket mode works, the dual-mode afterburner E2 is not additionally supplied with kerosene for combustion; the high-temperature and high-pressure airflow flowing out of the dual-mode afterburner E2 is finally discharged at a high speed through the tail nozzle E3 with a geometrically adjustable convergent-divergent structure to generate thrust.

A method of varying a crown turbine-based engine thrust with flight mach number for a variable cycle turbine rocket engine using a crown turbine system, characterized by:

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

secondly, 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 air turbine rocket mode is switched to the hybrid exhaust turbine fan mode, and the specific process is as follows:

1) The engine is firstly gradually opened, and then the mode selection valve B1 is opened;

2) The rear variable-area duct adjusting mechanism D5 gradually expands outwards;

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

4) After the rotating speed of the high-pressure compressor B2 reaches an ignition critical value, the main combustion chamber B4 is ignited, meanwhile, the rocket gas generator system C is gradually throttled, and the low-pressure turbine A3 and the crown turbine A4 jointly drive the fan A1 to rotate;

5) After the rotating speed of the high-pressure compressor B2 reaches an independent working critical value, the rocket combustion chamber C1 is completely flamed out, the low-pressure turbine A3 independently drives the fan A1 to rotate, and the dual-mode afterburner E2 is completely flamed out;

6) And continuously adjusting the variable-area duct adjusting mechanism D5 to expand outwards until the rocket duct D4 is completely closed, and completing the mode conversion from the air turbine rocket mode to the mixed exhaust turbofan mode by the engine.

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

1) The engine is firstly gradually closed to a front mode selection valve B1;

2) The rear variable-area duct adjusting mechanism D5 gradually retracts to open a rocket duct D4;

3) The rocket combustion chamber C1 starts to ignite and gradually generates rich combustion gas, and the coronally generated turbine A4 and the low-pressure turbine (A3) simultaneously drive the fan A1; the dual-mode afterburner E2 gradually reduces the supply amount of kerosene to control the total residual oxygen coefficient of the engine;

4) After the mass flow of the rich fuel gas generated by the rocket combustion chamber C1 reaches a critical value, the kerosene supply of the main combustion chamber B4 and the dual-mode afterburner E2 is cut off;

5) And continuously synchronously adjusting the retraction of the front mode selection valve B1 and the rear variable-area duct adjusting mechanism D5 until the inner duct D3 is completely closed, gradually decelerating and stopping the conventional high-pressure core machine system B, and finishing the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode by the engine.

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 in 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 conventional gas turbine engine has obvious insufficient thrust, obviously reduced fuel efficiency and even cannot normally work to provide power in the Mach number range of 2.5-4.0 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. The invention adopts a coronal turbine structure, reduces the difficulty of realizing the parallel design structure of two different gas generators, keeps the compact axial layout of the engine and brings positive benefits for shortening the strength of the lifting shaft of the rotating shaft; meanwhile, a remarkable process control benefit is provided for the bidirectional mode conversion process between the two working modes of the engine, the low-pressure shaft rotor does not stop in the starting and closing processes of the two core machines, and a feasible basis is provided for the mode conversion with stable thrust.

3. The invention shares the low-pressure turbine scheme and the crown turbine scheme, 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 use requirement of wide-speed-range hypersonic flight 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 of the present canopy turbine system;

FIGS. 1-2 are partial sectional elevation views of a variable cycle turbine rocket engine of the inventive shrouded turbine system;

FIGS. 1-3 are partial sectional elevation views of a variable cycle turbine rocket engine of the present canopy turbine system;

FIG. 2-1 is a schematic view of a mixed exhaust turbofan configuration of a shrouded turbine system of the present invention;

2-2 are schematic views of an air turbine rocket mode of the present canopy turbine system;

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 shows the total inlet temperature of the low-pressure turbine of the system as a function of the Mach number of the flight.

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: 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. In each scheme of the invention, a design method of a mechanical series structure layout of two combustion chambers and matching impellers thereof is adopted. 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 the total pressure increase ratio of about 10-25 and the 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 can 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.

2. 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 bypass design method and an outer bypass design method which can simultaneously take the requirements of the two modes into consideration.

1) For a mixed exhaust turbofan mode of a crown turbine system, an inner duct is adopted to be combined with a front mode selection valve B1 and a rear variable area duct adjusting mechanism D5 to do synchronous action, so that the complete opening/closing control of the inner duct D3 where a conventional high-pressure core system B and a low-pressure turbine A3 are located is realized, and thus, a high-pressure compressor B2 in the inner duct D3 can pressurize airflow again, so that the total pressurization ratio of an engine is higher in the mixed exhaust turbofan mode, and meanwhile, the high-pressure compressor B2 has better air suction capacity, so that 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;

2) When the air turbine rocket mode of the coronene turbine system is adopted, the rear variable-area duct adjusting mechanism D5 is matched with the front mode selection valve B1 to adjust and close the inner duct D3 where the conventional high-pressure core machine system B and the low-pressure turbine A3 are located. The air flow is forced to flow into the outer duct D2 completely at the position D1, the total duct ratio of the engine reaches a larger state by controlling the rich-fuel gas flow of the rocket duct D4, and meanwhile, the high-pressure compressor B2 stops working due to the fact that the inner duct D3 is closed, the supercharging capacity of the fan A1 is lower, and therefore the requirement of low total supercharging ratio can be met.

3) For the air turbine rocket mode of the shrouded turbine system, the method of shrouding the turbine system A4 on the blade shroud of the low-pressure turbine A3 is adopted, so that the difficulty of parallel design of two different gas generators is reduced, the compact axial layout of the engine is kept, and positive benefits are brought to the shortening of the strength of the lifting shaft of the rotating shaft; meanwhile, a remarkable process control benefit is provided for the bidirectional mode conversion process between the two working modes of the engine, the low-pressure shaft rotor does not stop in the starting and closing processes of the two core machines, and a feasible basis is provided for the mode conversion with stable thrust.

The invention is further explained below with reference to the drawings:

a variable cycle turbine rocket engine using a coronal turbine system is shown in figures 1-1, 1-2 and 1-3, and is provided with two sets of mutually independent gas generator systems, wherein 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; when the conventional high-pressure core system B works, the engine works in a mixed exhaust turbofan mode; when the rocket gas generator system C is in operation, the engine operates in an air turbine rocket mode; when the flying Mach number is within the range of 0.0-2.5, the hybrid exhaust turbofan mode is used for working, and when the flying 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: as shown in fig. 1-2 and 1-3, the low-pressure rotor system a is provided with a low-pressure rotor turbine which is integrally manufactured but is divided into an inner duct and an outer duct, wherein the low-pressure rotor turbine is a crown turbine A4 in the outer rocket duct D4, and the low-pressure rotor turbine is a low-pressure turbine A3 in the inner duct D3; the shrouded turbine A4 movable blade is fixedly connected above the shroud of the low-pressure turbine A3 movable blade; when the gas drives the low-pressure turbine A3 or the coronally generated turbine A4, the low-pressure rotating shaft A2 and the fan A1 can be simultaneously rotated; when the engine works in the mixed exhaust turbofan mode, the oxygen-enriched gas at the outlet of the conventional core system B drives a low-pressure turbine A3 to rotate and bring a fan A1 to suck and compress air flow; when the engine works in the air turbine rocket mode, the rich combustion gas at the outlet of the rocket gas generator system C drives a crown turbine A4 to suck and compress air flow with a rotating fan A1; the gas flow and the air flow are mixed and secondarily ignited in the afterburning boosting system E, and are discharged at a high speed to generate thrust.

Supplementary notes 1

The corotron turbine scheme is as shown in figures 2-1 and 2-2, when a mixed exhaust turbofan mode is adopted, as shown in figure 2-1, a rocket combustion chamber C1 of a rocket gas generator system C is closed, and a main combustion chamber B4 of a conventional high-pressure core engine system B is opened; 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 is composed of a fan A1, a low pressure turbine A3, a shrouded turbine A4, and a low pressure rotating shaft A2; the shrouded turbine A4 movable blade is fixedly connected above a low-pressure turbine A3 movable blade shroud, and the low-pressure turbine A3 movable blade and the shrouded turbine A4 movable blade rotate at the same constant angular speed as a whole; a labyrinth sealing structure is arranged on the edge of a movable blade crown of the low-pressure turbine A3, so that the airflow flow of an inner duct D3 where the low-pressure turbine is located and the airflow flow of a rocket duct D4 where the shrouded turbine is located are independent of each other; when the mixed exhaust turbine fan mode is adopted for working, oxygen-enriched gas flow generated by combustion of the main combustion chamber B4 blows the low-pressure turbine A3 through the high-pressure turbine B5, and the low-pressure turbine A3 rotates and simultaneously drives the crown turbine A4, the low-pressure rotating shaft A2 and the fan A1 to rotate; when the air turbine rocket mode is adopted, the crown turbine A4 is blown by the rich combustion gas flow generated by the combustion of the rocket combustion chamber C1, and the low-pressure turbine A3, the low-pressure rotating shaft A2 and the fan A1 are simultaneously rotated by the crown turbine A4.

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; the main combustion chamber B4 combusts to generate oxygen-enriched 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 partial air flow flowing through the fan A1 is sucked and 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 inner duct inlet, and the front mode selection valve B1 and a rear variable-area duct adjusting mechanism D5 synchronously act to realize the control of completely opening/closing the inner duct D3 where the conventional high-pressure core machine system B and the low-pressure turbine A3 are located.

Further, as shown in fig. 1-2, the rocket gas generator system C includes a rocket combustion chamber C1, a rocket nozzle C2; the rocket combustion chamber C1 combusts to generate rich fuel gas, the rich fuel gas flows into the crown turbine A4 to expand and do work, and the crown turbine A4 drives the low-pressure shaft to rotate A2, so that the fan A1 pressurizes air flow sucked into the engine.

Further, as shown in fig. 1-3, the bypass flow system D includes a high-pressure compressor inlet front bypass ring D1, an outer bypass D2, an inner bypass D3, a rocket bypass D4, and a rear variable-area bypass adjusting mechanism D5; the outer duct D2 and the inner duct D3 are formed by splitting at the front splitter ring D1 of the inlet of the high-pressure compressor; the air flow passing through the outer duct D2 finally flows directly into the mixer E1; the airflow passing through the inner duct D3 finally flows into the mixer E1 through the rear variable-area duct adjusting mechanism D5; the rocket duct D4 is a flow passage where rich-combustion gas flow generated by the rocket gas generator system C is located, and the rich-combustion gas flow flows into the mixer E1 through the rear variable-area duct adjusting mechanism D5 after flowing through the crown turbine A4.

Further, as shown in fig. 1-3, the rear variable-area duct adjusting mechanism D5 is capable of controlling the air flow of the inner duct D3 and the rocket duct D4 simultaneously; when the mixed exhaust turbine fan mode is used for working, the rear variable-area duct adjusting mechanism D5 adjusts and closes the rocket duct D4 where the coronal turbine A4 is located; when the air turbine rocket mode is used, the rear variable-area duct adjusting mechanism D5 is matched with the front mode selection valve B1 to adjust and close the inner duct D3 where the conventional high-pressure core machine system B and the low-pressure turbine A3 are located.

Further, as shown in fig. 1-2, the main combustion chamber B4 employs an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-enriched combustion, so as to ensure that when the mixed exhaust turbofan mode is used for working, the downstream dual-mode afterburner E2 has enough oxygen to perform afterburning.

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; when operating in the mixed exhaust turbofan mode, the mixer E1 is used for mixing the air flow flowing in the outer duct D2 and the oxygen-enriched gas flow flowing in the inner duct D3; when operating in the air turbine rocket mode, the mixer E1 is used for mixing the air flow flowing in from the outer duct D2 and the rich gas flow flowing in from the rocket duct D4; the airflow mixed by the mixer E1 enters a dual-mode afterburner E2 for afterburning combustion; when the mixed exhaust turbine fan mode works, the dual-mode afterburner E2 can freely adjust to carry out afterburning or not by controlling the supply amount of kerosene; when the air turbine rocket mode works, the dual-mode afterburner E2 is not additionally supplied with kerosene for combustion; the high-temperature and high-pressure airflow flowing out of the dual-mode afterburner E2 is finally discharged at a high speed through the tail nozzle E3 with a geometrically adjustable convergent-divergent structure to generate thrust.

A method for changing engine thrust along with flight Mach number based on a crown turbine is characterized in that:

4) After the rotating speed of the high-pressure compressor B2 reaches an ignition critical value, the main combustion chamber B4 is ignited, meanwhile, the rocket gas generator system C is gradually throttled, and the low-pressure turbine A3 and the crown turbine A4 drive the fan A1 to rotate together;

1) The engine is firstly gradually closed to a front mode selection valve B1;

Further, when the engine operates in a mixed exhaust turbofan mode within a flight Mach number range of 0.0 to 2.5, thrust and specific impulse characteristics reach set values within a speed range of 0.0 to 2.5 Mach numbers; 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.

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 more than 2,5, the inlet flow rate of the rocket chamber is not affected by the airflow and is only related to the combustion materials, so when the Mach number is more than 2,5, the inlet flow rate shows 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.

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 variable cycle turbine rocket engine using a coronal turbine system 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), one branch duct flowing 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 comprise: a hybrid exhaust turbofan mode and an air turbine rocket mode; when the conventional high-pressure core system (B) is operated, the engine is operated in a mixed exhaust turbofan mode; when the rocket gas generator system (C) is working, the engine works in an air turbine rocket mode; when the flying Mach number is within the range of 0.0-2.5, the hybrid exhaust turbofan mode is used for working, and when the flying 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: the low-pressure rotor system (A) is provided with a low-pressure rotor turbine which is integrally manufactured but is respectively arranged in an inner duct and an outer duct, wherein the low-pressure rotor turbine is positioned in an outer rocket duct (D4) and is a crown turbine (A4), and the low-pressure rotor turbine is positioned in an inner duct (D3) and is a low-pressure turbine (A3); the moving blade of the shrouded turbine (A4) is fixedly connected above the shroud of the moving blade of the low-pressure turbine (A3); when the gas drives the low-pressure turbine (A3) or the crown turbine (A4), the low-pressure rotating shaft (A2) and the fan (A1) can be simultaneously driven to rotate; when the engine operates in the mixed exhaust turbofan mode, the oxygen-enriched gas at the outlet of the conventional core system (B) drives a low-pressure turbine (A3) to take the rotating fan (A1) to suck and compress air flow; when the engine operates in said air turbine rocket mode, the rocket gas generator system (C) outlet rich gas drives the crown turbine (A4) with rotating fan (A1) to take in and compress the air flow; the gas flow and the air flow are mixed and secondarily ignited in the afterburning boosting system (E) and are discharged at high speed to generate thrust.

2. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 1 wherein: the low-pressure rotor system consists of a fan (A1), a low-pressure turbine (A3), a crown turbine (A4) and a low-pressure rotating shaft (A2); the movable blade shroud of the shrouded turbine (A4) is fixedly connected above the movable blade shroud of the low-pressure turbine (A3), and the movable blade shroud of the low-pressure turbine (A3) and the movable blade shroud of the shrouded turbine (A4) rotate at the same equal angular speed as a whole; a labyrinth sealing structure is arranged at the edge of a movable blade crown of the low-pressure turbine (A3), so that the airflow of an inner duct (D3) where the low-pressure turbine is located and the airflow of a rocket duct (D4) where the shrouded turbine is located are independent; when the mixed exhaust turbine fan mode is adopted for working, oxygen-enriched gas flow generated by combustion of the main combustion chamber (B4) blows the low-pressure turbine (A3) through the high-pressure turbine (B5), and the low-pressure turbine (A3) rotates and simultaneously drives the crown-rotating turbine (A4), the low-pressure rotating shaft (A2) and the fan (A1); when the air turbine rocket mode is adopted, the fuel-rich gas flow generated by the combustion of the rocket combustion chamber (C1) blows the crown turbine (A4), and the crown turbine (A4) rotates and simultaneously rotates the low-pressure turbine (A3), the low-pressure rotating shaft (A2) and the fan (A1).

3. A variable cycle turbo-rocket engine using a shrouded turbine system 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); the main combustion chamber (B4) combusts to generate oxygen-enriched 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 partial air flow flowing through the fan (A1) is sucked in and is 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 inner duct inlet, and the front mode selection valve (B1) can synchronously act with a rear variable-area duct adjusting mechanism (D5) to realize the control of completely opening/closing the inner duct (D3) where the conventional high-pressure core machine system (B) and the low-pressure turbine (A3) are located.

4. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 1 wherein: the rocket gas generator system (C) comprises a rocket combustion chamber (C1) and a rocket nozzle (C2); the rocket combustion chamber (C1) burns to generate rich fuel gas, the rich fuel gas flows into the crown turbine (A4) to expand and do work, and the crown turbine (A4) drives the low-pressure shaft to rotate (A2) so that the fan (A1) pressurizes the air flow sucked into the engine.

5. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 1 wherein: the bypass flow system (D) comprises a high-pressure compressor inlet front bypass ring (D1), an outer bypass (D2), an inner bypass (D3), a rocket bypass (D4) and a rear variable-area bypass adjusting mechanism (D5); the outer duct (D2) and the inner duct (D3) are formed by splitting at the position of a splitter ring (D1) in front of an inlet of the high-pressure compressor; the air flow passing through the bypass (D2) finally flows directly into the mixer (E1); the air flow passing through the inner duct (D3) finally flows into the mixer (E1) through the rear variable-area duct adjusting mechanism (D5); the rocket duct (D4) is a flow channel where rich fuel gas flow generated by the rocket gas generator system (C) is located, and the rich fuel gas flow flows into the mixer (E1) through the rear variable-area duct adjusting mechanism (D5) after flowing through the crown turbine (A4).

6. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 5 wherein: the rear variable-area duct adjusting mechanism (D5) can simultaneously control the airflow of the inner duct (D3) and the rocket duct (D4); when the mixed exhaust turbine fan mode is used for working, the rear variable-area duct adjusting mechanism (D5) adjusts and closes the rocket duct (D4) where the crown turbine (A4) is located; when the air turbine rocket mode is used, the rear variable-area duct adjusting mechanism (D5) is matched with the front mode selection valve (B1) to adjust and close the inner duct (D3) where the conventional high-pressure core machine system (B) and the low-pressure turbine (A3) are located.

7. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 4 wherein: the main combustion chamber (B4) adopts an air-kerosene constant pressure combustion system, and the main combustion chamber is oxygen-enriched combustion so as to ensure that the downstream dual-mode afterburning chamber (E2) has enough oxygen to carry out afterburning when the mixed exhaust turbine fan works in a mode.

8. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 4 wherein: the rocket combustion chamber (C1) adopts a liquid oxygen-kerosene constant pressure combustion system, in order to control the total temperature of the outlet of the rocket combustion chamber (C1), the rocket combustion chamber (C1) adopts rich combustion, and meanwhile, when the air turbine rocket mode is used for working, the downstream dual-mode afterburner (E2) can finish secondary combustion without additionally supplying kerosene.

9. A variable cycle turbo-rocket engine using a shrouded turbine system according to claim 1 wherein: the afterburning boosting system (E) comprises a mixer (E1), a dual-mode afterburner (E2) and a tail nozzle (E3); when operating in the mixed exhaust turbofan mode, the mixer (E1) is used for mixing the air flow flowing in the outer duct (D2) and the oxygen-enriched gas flow flowing in the inner duct (D3); -said mixer (E1) is intended to mix the air flow coming in from the bypass (D2) with the rich gas flow coming in from the rocket bypass (D4) when operating in said air turbine rocket mode; the air flow mixed by the mixer (E1) enters a dual-mode afterburner (E2) for afterburning combustion; when the mixed exhaust turbine fan mode works, the dual-mode afterburner (E2) can freely adjust the afterburning or not by controlling the supply amount of kerosene; when the air turbine rocket mode works, the dual-mode afterburner (E2) is not additionally supplied with kerosene for combustion; the high-temperature and high-pressure airflow flowing out of the dual-mode afterburner (E2) is finally discharged at high speed through a tail nozzle (E3) with a geometrically adjustable convergent-divergent structure to generate thrust.

10. A method of varying a thrust of a variable cycle turborocket engine based on a shrouded turbine system as a function of flight mach number according to any of claims 1 to 9, wherein:

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

11. A method of varying engine thrust as a function of flight mach number based on a shrouded turbine in accordance with claim 10, wherein: the method is characterized in that the air turbine rocket mode is switched to the mixed exhaust turbine fan mode, and the specific process is as follows:

1) The engine gradually opens a front mode selection valve (B1);

2) The rear variable-area duct adjusting mechanism (D5) gradually expands outwards;

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

4) After the rotating speed of the high-pressure compressor (B2) reaches an ignition critical value, the main combustion chamber (B4) is ignited, meanwhile, the rocket gas generator system (C) is gradually throttled, and the low-pressure turbine (A3) and the crown turbine (A4) jointly drive the fan (A1) to rotate;

5) After the rotating speed of the high-pressure compressor (B2) reaches an independent working critical value, the rocket combustion chamber (C1) is completely flamed out, the low-pressure turbine (A3) independently drives the fan (A1) to rotate, and the dual-mode afterburner (E2) is completely flamed out;

6) And continuously adjusting the variable-area duct adjusting mechanism (D5) to expand until the rocket duct (D4) is completely closed, and finally completing the mode conversion from the air turbine rocket mode to the mixed exhaust turbofan mode by the engine.

12. A method of varying engine thrust as a function of flight mach number based on a shrouded turbine in accordance with claim 10, wherein: the mode is switched from the mixed exhaust turbofan mode to the air turbine rocket mode, and the specific process is as follows:

1) The engine is gradually closed to a front mode selection valve (B1);

2) The back variable-area duct adjusting mechanism (D5) gradually retracts to open the rocket duct (D4);

3) The rocket combustion chamber (C1) starts to ignite and gradually generates rich combustion gas, and the crown turbine (A4) and the low-pressure turbine (A3) drive the fan (A1) simultaneously; the dual-mode afterburner (E2) gradually reduces the supply amount of kerosene to control the total residual oxygen coefficient of the engine;

4) Cutting off kerosene supply of a main combustion chamber (B4) and a dual-mode afterburning chamber (E2) after the mass flow of rich fuel gas generated by the rocket combustion chamber (C1) reaches a critical value;

5) And continuously synchronously adjusting the front mode selection valve (B1) and the rear variable-area duct adjusting mechanism (D5) to be adducted until the inner duct (D3) is completely closed, gradually decelerating and stopping the conventional high-pressure core machine system (B), and finishing the mode conversion from the mixed exhaust turbofan mode to the air turbine rocket mode by the engine.

13. The method of claim 11, wherein the method further comprises varying a thrust of the shrouded turbine based engine as a function of a flight mach number: 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.