CN111594315B

CN111594315B - Composite mechanism full-flow circulation supersonic propulsion system and working method thereof

Info

Publication number: CN111594315B
Application number: CN202010288416.3A
Authority: CN
Inventors: 郑耀; 黄日鑫; 王浩添; 崔涛
Original assignee: Zhejiang Institute Of Turbomachinery And Propulsion System
Current assignee: Zhejiang Institute Of Turbomachinery And Propulsion System
Priority date: 2020-04-07
Filing date: 2020-04-14
Publication date: 2021-03-05
Anticipated expiration: 2040-04-14
Also published as: CN111594315A

Abstract

The invention provides a composite mechanism full-flow circulation supersonic propulsion system and a working method thereof, wherein the composite mechanism full-flow circulation supersonic propulsion system comprises an air inlet channel, an air compressor, a diffusion duct, a shaft, an open type pre-combustion driver, a closed type pre-combustion driver, a turbine, a afterburning thruster and a tail nozzle; the air inlet channel, the air compressor, the shaft, the turbine, the afterburning thruster and the tail nozzle are sequentially connected; the compressor is connected with the turbine through a shaft; the turbine is positioned on the air inlet side of the afterburning thruster, and the tail nozzle is communicated with the outlet of the afterburning thruster; the open type pre-combustion driver and the closed type pre-combustion driver are both positioned behind the gas compressor, and the outlets of the open type pre-combustion driver and the closed type pre-combustion driver are opposite to the turbine; the open type pre-combustion drivers and the closed type pre-combustion drivers are alternately distributed in the diffusion duct in the annular space outside the shaft. The system gives full play to the propulsion advantages of all subsystems in the aspects of specific impulse and thrust-weight ratio as much as possible, thereby realizing the supersonic speed combined propulsion system which has the advantages of quick self-starting, multiple degrees of freedom, compound working mechanism, high efficiency and economy.

Description

Composite mechanism full-flow circulation supersonic propulsion system and working method thereof

Technical Field

The invention relates to the technical field of supersonic propulsion, in particular to a composite mechanism full-flow circulation supersonic propulsion system and a working method thereof.

Background

In recent years, with the progress of science and technology and the traction of military requirements, the adjacent space gradually draws attention from various countries due to the potential military and civil values of the adjacent space, and a series of researches are carried out on the adjacent space by the countries. The near space aircraft propulsion system not only requires the ability of freely entering the near space, but also requires the ability of high-altitude high-speed cruising operation, and meanwhile, the near space aircraft propulsion system has the characteristics of high performance, light weight, high reliability, low cost, long range, no boosting, reusability, wide range of adaptive speed and the like.

The existing propulsion systems, including various aero-engines, ramjets, rocket engines, and piston engines, all have their own performance advantages and ideal flight airspace. Aircraft engines are characterized by high performance, but are difficult to use in propulsion systems for aircraft with speeds above 20km or speeds greater than mach 3.0; the ramjet has higher performance and flight Mach number, but has poorer flight mobility, and the starting speed problem needs to be solved by a booster; the rocket engine is not limited by the height and initial speed, but has low performance and high propellant consumption rate.

By adopting a power circulation mechanism composite mode, if an air-breathing propulsion system is used under low flight Mach number, and a rocket engine is switched to work when the flight Mach number is higher, the maneuverability and flexibility of the air-craft can be improved inevitably, the cost of entering space can be greatly reduced, and therefore, the development of air-breathing combined power is significant. However, the combined engine which is currently researched, namely a TBCC engine or an RBCC engine, has the problem of discontinuous thrust in the conversion process of working modes of different cycles; moreover, the structure of the mode conversion device is relatively complex, resulting in an increase in the weight of the combined engine, which brings about a number of drawbacks such as a reduction in the payload of the aircraft and an increase in the fuel consumption.

Disclosure of Invention

Aiming at the defects of the prior art, the invention aims to provide an efficient and reasonable turbojet and rocket combined propulsion system and a working method thereof, which can give full play to the propulsion advantages of each subsystem in the aspects of specific impulse and thrust-weight ratio as much as possible, thereby realizing an ultrasonic combined propulsion system with multiple degrees of freedom, a composite working mechanism, high efficiency and economy.

In order to achieve the purpose, the invention provides a composite mechanism full-flow circulation supersonic propulsion system, which comprises an air inlet channel, an air compressor, a diffusion duct, a shaft, an open type pre-combustion driver, a closed type pre-combustion driver, a turbine, a afterburning thruster and a tail nozzle; the air inlet passage, the air compressor, the shaft, the turbine, the afterburning thruster and the tail nozzle are sequentially connected, the air compressor is connected with the turbine through the shaft, the turbine is positioned on the air inlet side of the afterburning thruster, and the tail nozzle is communicated with the outlet of the afterburning thruster; the opening type pre-combustion driver and the closed type pre-combustion driver are positioned on the side of an air outlet of the air compressor, and outlets of the opening type pre-combustion driver and the closed type pre-combustion driver are opposite to the turbine and used for pushing the turbine; the open type pre-combustion driver and the closed type pre-combustion driver are alternately and circumferentially distributed in the diffusion duct in the annular space outside the shaft.

Preferably, the working medium of the open pre-combustion driver is fuel and air, and the fuel can be any hydrocarbon fuel in principle.

Preferably, a plurality of specially designed air holes are arranged on the circumferential direction of the wall surface of the open type pre-combustion driver, so that the working medium and the air can be conveniently mixed.

Preferably, the closed pre-combustion driver takes self-contained propellant as working medium, and the propellant is monopropellant, bipropellant or tripropellant. The propellant with the self-contained function is universal with the traditional rocket engine propellant, for example, a single-component propellant, a double-component propellant, a three-component propellant and the like can be selected according to task needs, and the single-component propellant can be selected from hydrogen peroxide, single push-3, anhydrous hydrazine, methylhydrazine, ADN, HAN and other novel green propellants and the like; the two-component oxidant can be liquid oxygen, hydrogen peroxide, dinitrogen tetroxide, nitrous oxide and the like, and the fuel can be various hydrocarbon fuels. The selection of the working medium of the open type pre-combustion driver and the closed type pre-combustion driver can be determined according to the performance of the working medium according to the task type.

Preferably, the open type pre-combustion driver and the closed type pre-combustion driver are opposite to the turbine and can independently push the turbine.

Preferably, the open type pre-combustion drivers and the closed type pre-combustion drivers are independent cylindrical structures, a plurality of open type pre-combustion drivers and closed type pre-combustion drivers are alternately arranged in the circumferential direction of the diffuser duct, the arrangement number is not limited in principle, and the arrangement number can be flexibly selected according to the thrust and the size scale of the engine.

Preferably, the inlet end of the diffuser duct has a channel with an annular cross section with an area gradually increasing along the air flow direction, and the outlet end of the diffuser duct has a channel with an annular cross section with an area gradually decreasing along the air flow direction.

Preferably, the compressor is used for sucking air from the atmospheric environment and supercharging, a supercharging region capable of accommodating the compressor is enclosed between the air inlet channel and the shaft, and the area of the annular section of the supercharging region is gradually reduced along the air circulation direction.

Preferably, the air outlet end of the tail pipe is provided with a channel with a circular cross section along the air circulation direction and with gradually reduced area.

The invention also provides a working method of the composite mechanism full-flow circulation supersonic propulsion system, which specifically comprises the following working modes:

a rocket starting mechanism: under the mechanism, the system mainly drives a turbine by the work of a closed pre-combustion driver, directly drives a gas compressor to work, and can finish starting in a very short time along with the rapid pressure build-up of a afterburning thruster without any auxiliary system depending on the traditional starting process; the starting mechanism working mode can quickly realize the self-starting of the propulsion system;

complete turbojet mechanism: under the complete turbojet mechanism, the propulsion system drives the turbine completely by the open type pre-combustion driver, so that high specific impulse characteristic and fuel economy are realized, and the voyage can be greatly improved; the complete turbojet mechanism is suitable for a low-speed cruise task;

the full mechanism working mode is as follows: all open type pre-combustion drivers and all closed type pre-combustion drivers work on the turbine, so that the maximum power work of the engine can be realized, and the turbine works at full load; at the moment, the thrust is maximum, and the fuel economy of achieving the target flight speed is optimal; the full-mechanism working mode is suitable for a takeoff acceleration climbing stage;

a composite working mechanism: the open type pre-combustion driver and the closed type pre-combustion driver are combined in any number; the composite working mechanism can fully combine the respective advantages of the turbojet and the rocket engine, realize the adjustment of different thrust levels, and is favorable for optimizing and realizing the optimal specific impulse and fuel economy under different flight Mach numbers; the compound operating mechanism mode is suitable for high-speed cruising and tactical maneuvers.

The invention has the beneficial effects that:

(1) the two circulation mechanisms of the turbojet and the rocket are mutually independent, and any number of working combinations can be realized. The invention relates to a turbojet and rocket combined supersonic propulsion system which is provided with two types of pre-combustion drivers, wherein an open type (open type pre-combustion driver) medium is fuel and air, a closed type (closed type pre-combustion driver) medium is a general rocket propellant, the working processes of a turbojet engine and a rocket engine are respectively realized, the working states of the two types of pre-combustion drivers are not influenced mutually, and any pre-combustion driver can be opened and closed according to actual needs to generate different circulation and thrust combinations. Energy-containing combustion products generated by the pre-combustion driver push the turbine to do work, and the compressor is driven to compress air to flow. And the high-temperature and high-pressure gas flowing out from the rear of the turbine continues to be mixed and combusted with the residual air in the afterburning thruster, and thrust is generated through the tail nozzle. When the closed type driver is completely opened, the problem of thrust trap caused by the fact that the traditional aircraft engine is limited to the use limit of the temperature in front of the turbine can be solved. When the open type driver is completely opened and the closed type driver is completely closed, a pure aviation engine working mechanism can be realized, higher specific impulse performance is realized, and more economic fuel consumption rate is obtained.

(2) All the working media are in full flow to do work on the turbine, and the circulation efficiency is high. The working medium of the pre-combustion driver flows through the turbine in a full-flow mode to do work, and the high-temperature energetic medium of the closed driver and the large-flow characteristic of the open pre-combustion driver are fully utilized, so that the work capacity of the turbine is far higher than that of the traditional aero-engine and air turbine rocket engine.

(3) Under the same task section, the compressor and the turbine can realize single-stage design, and the compressor and the turbine have simple structure, light weight and reliability. Since both types of pre-combustion drivers are of independent cylindrical configuration and can be alternately arranged circumferentially in front of the turbines, a common use of the same stage of turbines is achieved. Compared with an air turbine rocket engine, the multi-stage turbine pre-combustion engine has the advantages that the mutual influence among the multi-stage turbines is reduced, more importantly, the structure and the quality of the multi-stage turbines are reduced, a better multifunctional integration effect is obtained, and the two pre-combustion drivers adopt the full-flow post-combustion characteristic, so that the turbine efficiency is greatly improved.

According to the invention, through reasonable matching of the two independent pre-combustion drivers, the energy efficiency of the propulsion system is improved and the fuel economy is improved under the condition of minimum propellant consumption. Moreover, the cooperation of the two pre-combustion drivers can prevent the condition that any driver stops in the air or cannot be started automatically, and greatly improves the safety performance of the aircraft; meanwhile, the self-starting system can completely realize the self-starting of the propulsion system, and solves the problem that the existing turbine aviation propulsion system cannot be self-started. The composite working mechanism can enrich the task types of the engine and expand the application range of the engine to the non-designed flight envelope.

The features and advantages of the present invention will be described in detail by embodiments in conjunction with the accompanying drawings.

Drawings

FIG. 1 is a schematic diagram of the internal structure of a compound mechanism full flow circulation supersonic propulsion system of the present invention;

FIG. 2 is a cross-sectional view of an arrangement of an open pre-combustion driver and a closed pre-combustion driver of a combined-mechanism full-flow-circulation supersonic propulsion system according to the present invention;

wherein, 1 is an air inlet channel, 2 is a gas compressor, 3 is a diffusion duct, 4 is an axis, 5 is an open type pre-combustion driver, 6 is a closed type pre-combustion driver, 7 is a turbine, 8 is a afterburning thruster, and 9 is a tail nozzle.

Detailed Description

Referring to fig. 1, the invention provides a compound mechanism full flow circulation supersonic propulsion system, which comprises an air inlet 1, an air compressor 2, a diffusion duct 3, a shaft 4, an open type pre-combustion driver 5, a closed type pre-combustion driver 6, a turbine 7, a afterburning thruster 8 and a tail nozzle 9. The air inlet channel 1, the air compressor 2, the shaft 4, the turbine 7, the afterburning thruster 8 and the tail nozzle 9 are connected in sequence. The compressor 2 is connected with the turbine 7 through the shaft 4, and the gas generated by the open type pre-combustion driver 5 and the closed type pre-combustion driver 6 pushes the turbine 7 to drive the compressor 2 to do work on the incoming air flow. The compressor 2 has the functions of sucking air from the atmospheric environment and boosting pressure, the design is full-flow afterburning, and the air completely enters the afterburning thruster 8 after passing through the diffusion duct 3. The turbine 7 is positioned on the air inlet side of the afterburning thruster 8, the open type pre-combustion driver 5 and the closed type pre-combustion driver 6 share the same turbine 7, and the tail nozzle 9 is communicated with the outlet of the afterburning thruster 8. The afterburning thruster 8 is used for efficiently mixing and combusting high-temperature gas and high-pressure air, and expanding and doing work in the tail nozzle 9 to generate thrust. The system comprises two pre-combustion drivers, wherein an open type pre-combustion driver 5 and a closed type pre-combustion driver 6 are positioned behind the gas compressor 2, and outlets of the open type pre-combustion driver and the closed type pre-combustion driver are opposite to the turbine 7 and can independently push the turbine 7.

Further, as shown in fig. 2, the open-type pre-combustion drivers 5 and the closed-type pre-combustion drivers 6 are independent cylindrical structures, alternately arranged circumferentially in the diffuser duct 3 outside the shaft 4. The outer wall of the open type pre-combustion driver 5 is provided with air holes to facilitate the mixing of the medium and the air. The closed pre-combustion driver 6 takes propellant as working medium, and the propellant is single-component propellant, double-component propellant or three-component propellant.

when a rocket starting mechanism works, the combined system is mainly powered by a closed pre-combustion driver 6, the closed pre-combustion driver 6 works to generate high-temperature and high-pressure rich combustion gas and air pressurized by an air compressor 2 simultaneously and independently apply work to a driving turbine 7 to drive the air compressor 2 to rotate, the turbine 7 after applying work exhausts air, the air and the rich combustion gas are mixed and combusted in a afterburning thruster 8, and the generated high-temperature gas is exhausted through a tail nozzle 9 to generate thrust. The working mode can be suitable for the takeoff phase, and the propulsion system is self-started.

When the full turbojet mechanism works, the open type pre-combustion driver 5 works, the closed type pre-combustion driver 6 is closed, air flows through the gas compressor 2 from the air inlet passage 1 to form air inflow with certain pressure and temperature, the pressurized air is mixed and combusted with fuel sprayed by the open type pre-combustion driver 5 through holes in the outer wall of the open type pre-combustion driver 5 to generate high-temperature and high-pressure gas, the high-temperature and high-pressure gas flows through a turbine 7 which is coaxial with the gas compressor 2 and 4, and partial internal energy of the gas is expanded in the turbine 7 to be converted into mechanical energy to drive the gas compressor 2 to rotate. The high-temperature lean fuel gas flowing out of the turbine 7 and the fuel sprayed secondarily in the afterburning thruster 8 are mixed and combusted with the incoming air flow in the afterburning thruster 8, and the expansion is accelerated through the tail nozzle 9 to do work to generate thrust. The full turbojet mechanism is suitable for low-speed cruising tasks.

When operating in the fully machined mode, both the open predator 5 and the closed predator 6 are fully operational. The closed pre-combustion driver 6 generates high-temperature and high-pressure rich combustion gas and drives the turbine 7 to do work, the gas enters the afterburning thruster 8 after doing work, the gas and the high-temperature lean-oil gas generated by the open pre-combustion driver 5 or fuel sprayed in the afterburning thruster 8 are mixed and combusted with air incoming flow in the afterburning thruster 8, and the generated high-temperature gas is discharged through the tail nozzle 9 to generate larger thrust. At the moment, all the pre-combustion drivers are fully opened to apply work to the turbine, the maximum power work of the engine can be realized, the full load work and the maximum thrust of the turbine are realized, and the fuel economy of achieving the target flight speed is optimal. The full-mechanism working mode is suitable for a takeoff acceleration climbing phase.

When working with a compound working regime, i.e. two types of pre-combustion drivers work in any number in combination. And closing the pre-combustion drivers of the designated types or the designated number according to the requirement of the thrust magnitude, and pushing the turbine 7 to do work by energy-containing combustion products generated by the rest pre-combustion drivers so as to drive the compressor 2 to compress air to flow. The high-temperature and high-pressure gas flowing out of the turbine 7 continues to be mixed and combusted with the residual air in the afterburning thruster 8, and thrust is generated through the tail nozzle 9. The composite working mechanism can fully combine the respective advantages of the turbojet and the rocket engine, realize the adjustment of different thrust levels, and is favorable for optimizing and realizing the optimal specific impulse and fuel economy under different flight Mach numbers. The compound operating mechanism mode is suitable for high-speed cruising and tactical maneuvers. In the mode, if all the closed type pre-combustion drivers work, the problem of thrust trap caused by the limitation of the traditional aircraft engine to the use limit of the temperature in front of the turbine can be solved.

The above embodiments are illustrative of the present invention, and are not intended to limit the present invention, and any simple modifications of the present invention are within the scope of the present invention.

Claims

1. The utility model provides a compound mechanism full flow circulation supersonic velocity propulsion system which characterized in that: the device comprises an air inlet passage (1), a gas compressor (2), a diffusion duct (3), a shaft (4), an open type pre-combustion driver (5), a closed type pre-combustion driver (6), a turbine (7), a afterburning thruster (8) and a tail nozzle (9), wherein the air inlet passage (1), the gas compressor (2), the shaft (4), the turbine (7), the afterburning thruster (8) and the tail nozzle (9) are sequentially connected, the gas compressor (2) is connected with the turbine (7) through the shaft (4), the turbine (7) is positioned on the air inlet side of the afterburning thruster (8), and the tail nozzle (9) is communicated with the outlet of the afterburning thruster (8); the open type pre-combustion driver (5) and the closed type pre-combustion driver (6) are positioned on the air outlet side of the air compressor (2), and the outlets of the open type pre-combustion driver (5) and the closed type pre-combustion driver (6) are opposite to the turbine (7) and used for pushing the turbine (7); the open type pre-combustion driver (5) and the closed type pre-combustion driver (6) are alternately and circumferentially distributed in the diffusion duct (3) in the annular space outside the shaft (4); the working medium of the open type pre-combustion driver (5) is fuel and air; the closed type pre-combustion driver (6) takes a propellant with itself as a working medium, and the propellant is a single-component propellant, a double-component propellant or a three-component propellant; the open type pre-combustion driver (5) and the closed type pre-combustion driver (6) are both over against the turbine (7) and can independently push the turbine (7);

the system specifically comprises the following working modes:

a rocket starting mechanism: under the mechanism, the system mainly drives a turbine (7) by the work of a closed pre-combustion driver (6) to directly drive a gas compressor (2) to work, and can complete starting in a very short time along with the rapid pressure build-up of a afterburning thruster (8) without any auxiliary system depending on the traditional navigation starting process; the starting mechanism working mode can quickly realize the self-starting of the propulsion system;

complete turbojet mechanism: under the complete turbojet mechanism, the propulsion system completely drives the turbine (7) by the open type pre-combustion driver (5), so that high specific impulse characteristic and fuel economy are realized, and the voyage can be greatly improved; the complete turbojet mechanism is suitable for a low-speed cruise task;

the full mechanism working mode is as follows: all open type pre-combustion drivers (5) and all closed type pre-combustion drivers (6) apply work to the turbine (7), so that the maximum power work of the engine can be realized, and the turbine (7) can work at full load; at the moment, the thrust is maximum, and the fuel economy of achieving the target flight speed is optimal; the full-mechanism working mode is suitable for a takeoff acceleration climbing stage;

a composite working mechanism: the open type pre-combustion driver (5) and the closed type pre-combustion driver (6) are combined to work in any number; the composite working mechanism can fully combine the respective advantages of the turbojet and the rocket engine, realize the adjustment of different thrust levels, and is favorable for optimizing and realizing the optimal specific impulse and fuel economy under different flight Mach numbers; the compound operating mechanism mode is suitable for high-speed cruising and tactical maneuvers.

2. The compound-mechanism full-flow-cycle supersonic propulsion system of claim 1, wherein: a plurality of air holes are circumferentially arranged on the wall surface of the open type pre-combustion driver (5) so as to facilitate the mixing of working media and air.

3. The compound-mechanism full-flow-cycle supersonic propulsion system of claim 1, wherein: the open type pre-combustion driver (5) and the closed type pre-combustion driver (6) are both independent cylindrical structures, and the plurality of open type pre-combustion drivers (5) and the closed type pre-combustion drivers (6) are alternately arranged in the circumferential direction of the diffuser duct (3).

4. The compound-mechanism full-flow-cycle supersonic propulsion system of claim 1, wherein: the air inlet end of the diffusion duct (3) is provided with a channel with the area of the annular section gradually increasing along the air circulation direction, and the air outlet end of the diffusion duct (3) is provided with a channel with the area of the annular section gradually decreasing along the air circulation direction.