RU2848345C1 - Hypersonic turbojet engine and its regulation method - Google Patents

Abstract

FIELD: aircraft engine building.

SUBSTANCE: hypersonic turbojet engine is proposed, in which a gas ejector with a cylindrical mixing chamber and a diffuser is installed between the inlet device and the compressor. The high-pressure channel of the gas ejector is connected on one side to the cavity located behind the turbine, and on the other side to the entrance to the mixing chamber, the low-pressure channel of the gas ejector is connected to the atmosphere on the one hand through the inlet device, and on the other hand to the entrance to the mixing chamber, at the outlet of the high-pressure channel of the gas ejector an adjustable nozzle is installed. Afterburner and hyperburst are used to boost thrust at super- and hypersonic flight speeds. The engine has a control program that ensures that the reduced rotor speed is maintained at a constant speed in the range of flight speeds from takeoff to four Mach numbers. The maximum flight speed is five Mach numbers.

EFFECT: invention makes it possible to exclude the influence of air temperature on the operation of the turbocharger in a wide range of flight speeds.

10 cl, 3 dwg

Description

The invention relates to aircraft engine manufacturing.

A turbojet engine (TRD) is known that contains a supersonic inlet device, a turbocompressor, an afterburner, and an outlet device (Theory and calculation of air-breathing engines. Ed. S.M. Shlyakhtenko. Moscow: Mashinostroenie, 1987, p. 16, Fig. 1.2).

A turbojet engine with ejector supercharging (TRD) is known, containing a gas ejector, the high-pressure channel of which is looped through a mixing chamber, a diffuser and a compressor, and the low-pressure channel is connected on one side to the atmosphere through an inlet device, and on the other side to the compressor through a mixing chamber and a diffuser (patent RU 2201518).

Methods for boosting a gas turbine engine are known, in particular, hyperboost - feeding water to the compressor inlet at flight speeds greater than three Mach numbers (Pismenny V.L. Methods for boosting a gas turbine engine. News of higher educational institutions. Mechanical Engineering, 2024, No. 9, pp. 120-130).

Methods for increasing the temperature of the gas in front of the turbine blades of a gas turbine engine are known, in particular, single-crystal blades and air-liquid cooling (Pismenny V.L. Methods and means of increasing the temperature of the gas in front of the turbine of a gas turbine engine. News of higher educational institutions. Mechanical Engineering, 2023, No. 6, pp. 108-118).

A stoichiometric turbojet engine is known in which the control law is used as the engine control program (Pismenny V.L. Concept of constructing a stoichiometric turbojet engine. News of higher educational institutions. Mechanical Engineering, 2023, No. 9, pp. 116-126).

A turbojet engine is known that uses a single-stage adjustable turbine (patent RU 2616137).

A method for regulating an axial compressor in a gas turbine engine system is known, which consists of feeding hot gas, taken from a channel located behind the turbine, into a channel located between the inlet device and the engine compressor (patent RU 2535186).

A method for increasing gas pressure is known, which consists of feeding a liquid under pressure of more than 5 MPa into the gas flow moving through a channel in the direction of gas movement (patent RU 2468260).

A method is known for reducing the air temperature at the inlet of a gas turbine engine using a heat exchange device through which fuel is supplied to the main combustion chamber of the engine (Certificate for Utility Model RU 29752).

The fundamental disadvantage of a turbojet engine, which limits the engine's speed capabilities, is the decrease in compressor performance with increasing flight speed (a decrease in the equivalent air flow rate).

The aim of the invention is to eliminate the said disadvantage.

The stated objective is achieved by the fact that a gas ejector with a mixing chamber and a diffuser is located between the inlet device and the compressor of a turbojet engine with an afterburner. The high-pressure channel of the gas ejector is connected on one side to a cavity located behind the turbine, and on the other side to the inlet of the mixing chamber. The low-pressure channel of the gas ejector is connected on one side to the atmosphere through the inlet device, and on the other side to the inlet of the mixing chamber. An adjustable nozzle is installed at the outlet of the high-pressure channel of the gas ejector. A water manifold with nozzles directed along the air flow is located in the mixing chamber. A heat exchanger is located in the diffuser, through which fuel is supplied to the main combustion chamber of the engine. The standard operating conditions of the engine are those corresponding to the conditions of the cruising flight mode of the aircraft. In this case, the relative reduced rotor speed is determined by the formula

Where - relative physical rotation frequency,

- physical rotation frequency,

- physical rotation frequency at cruise flight mode,

- braking temperature of the air at the compressor inlet in the cruising flight mode,

- the stagnation temperature of the air at the compressor inlet. The essence of the invention is that supercharging the compressor with hot, high-pressure gas using a gas ejector with an adjustable nozzle at the outlet of the high-pressure channel makes it possible to eliminate the influence of air temperature on turbocharger operation over a wide range of flight speeds, ensuring the implementation of the engine control law. (Theory and calculation of air-breathing engines. Ed. by S.M. Shlyakhtenko - M.: Mashinostroenie, 1987, p. 254), at which the air flow through the engine is the maximum possible.

Fig. 1 shows a diagram of a hypersonic turbojet engine (HTJE);

Fig. 2 shows the parameters of the gas turbine engine in take-off mode;

Fig. 3 shows the altitude-speed characteristics of the gas turbine engine.

The gas turbojet engine (Fig. 1) consists of an inlet device 1, a sonic gas ejector 2, a cylindrical mixing chamber 3, inside of which is a water manifold 4 with nozzles directed along the air flow, a diffuser 5, inside of which is a heat exchange device 6, through which fuel is supplied to the main combustion chamber of the engine, a turbocharger 7, an afterburner 8, inside of which is a fuel manifold 9, an outlet device (adjustable Laval nozzle) 10. Gas enters the high-pressure channel of the gas ejector from a cavity located behind the engine turbine, at the outlet of the high-pressure channel of the gas ejector an adjustable (with the help of a conical inner body) tapering nozzle 11 is installed. The low-pressure channel is connected to the atmosphere on one side through the inlet device, and on the other side - to the mixing chamber.

The engine operates as follows. At subsonic flight speeds up to Mach 0.9, the stagnation temperature of the air at the compressor inlet equal to the braking temperature of the outside air relative physical rotor speed respectively,

At flight speeds greater than M=0.9 and less than cruising speed (M~2.5), the stagnation temperature of the air at the compressor inlet equal to the air stagnation temperature at cruising flight mode - is maintained by supercharging the compressor with hot gas, the relative physical rotation frequency is equal to the relative reduced rotation frequency Fuel (kerosene) enters the main combustion chamber through heat exchanger 6. In the heat exchanger, the kerosene evaporates, absorbing some of the air's thermal energy. The air temperature at the compressor inlet decreases by approximately 6 degrees, and the calorific value of the kerosene increases by approximately 300 kJ/kg.

When cruising speed is reached, the pressure difference in the turbine increases to its maximum (due to the nozzle opening), the relative rotation frequency remains unchanged. the gas temperature in front of the turbine decreases.

At flight speeds exceeding cruising speed, the physical rotation frequency increases proportionally until the gas temperature limit in front of the turbine is reached , after which water is supplied to the mixing chamber under a pressure of more than 10 MPa (hyperforce) in an amount that ensures maintaining a constant air temperature at the outlet of the mixing chamber, which ensures that the following conditions are met:

After reaching the water flow limit, afterburner is activated, the relative reduced and relative physical rotation speeds decrease, and the engine gradually degenerates.

Below are the flight characteristics of the gas turbine engine (Fig. 1) with the initial data: takeoff thrust 20 tons; maximum gas temperature in front of the turbine 1900 K; afterburner temperature 2200 K; maximum temperature of the working blades 1350 K; air bleed coefficient for cooling the turbine (cooling intensity coefficient of the working blades 0.67); the degree of increase in compressor pressure under takeoff conditions turbine pressure reduction ratio under takeoff conditions Pressure losses in the inlet device, gas ejector, combustion chamber, and outlet device are standard; compressor and turbine efficiencies are standard. The ejector is sonic, the ratio of the nozzle outlet cross-sectional areas for the ejecting and ejected gases is α = 0.17. The engine midsection diameter is 1.7 m.

Fig. 2 shows the takeoff mode characteristics of the gas turbine engine. Fig. 3 shows the altitude-speed characteristics of the gas turbine engine (Fig. 1) for a typical flight trajectory of a hypersonic aircraft (here H is the flight altitude, R is the engine thrust, S is the specific fuel consumption, - gas temperature before turbine, - afterburner temperature; - temperature of the working blades, - air temperature at the compressor outlet, - relative rotor speed, - relative reduced rotor speed, - the degree of pressure increase in the compressor, - the degree of pressure reduction in the turbine, - overall engine efficiency; - relative fuel consumption; - relative water consumption; - excess air coefficient in the main combustion chamber; - excess air coefficient in the afterburner; - the share of gas extracted for compressor boosting.

The cruising speed is Mach 2.5. The ejector is turned on at Mach 0.9 and turned off at Mach 2.5. Up to Mach 4.0, the control law is implemented. At speeds M> 4.0 the engine is at the water flow limiter - the regulatory law is being implemented afterburner is turned on The maximum flight speed of the aircraft (M=5.0) is limited by the strength of the compressor blades Maximum engine fuel efficiency is achieved at flight speeds

The gas turbine engine (Fig. 1) has unique thrust and fuel consumption characteristics at sub- and hypersonic flight speeds (Fig. 2, Fig. 3): specific fuel consumption at takeoff is ~ 1.0 kg/(kgf⋅h), overall efficiency at a speed of four Mach is ~ 0.47, maximum flight speed is ~ 5M.

The gas turbine engine can be used as an engine for supersonic passenger and transport aircraft and as an engine for military (special) purpose aircraft.

For passenger and transport aircraft, the gas turbine engine is used without thrust augmentation: the maximum flight speed is Mach 3.3. Existing fifth-generation engine technologies (implemented in production) are sufficient to create such engines.

For military (special) purpose aircraft, gas turbine engines are used in combination with thrust augmentation devices (hyperforce, afterburner, see Fig. 1): maximum flight speed is 5M. To create such engines, it is necessary to develop new technologies that ensure an increase in the heat resistance of compressor blades to 1200 K, turbine blades - up to 1400 K, or the introduction of more effective methods of their cooling (protection) (Pismenny V.L. Methods and means of increasing the gas temperature in front of the turbine of a gas turbine engine. News of higher educational institutions. Mechanical Engineering, 2023, No. 6, pp. 108-118).

Claims (22)

1. A hypersonic turbojet engine comprising a supersonic inlet device, a turbocharger, an afterburner, an outlet device, characterized in that a gas ejector with a mixing chamber and a diffuser is located between the inlet device and the compressor, the high-pressure channel of the gas ejector is connected on one side to a cavity located behind the turbine, and on the other side to the inlet of the mixing chamber, the low-pressure channel of the gas ejector is connected on one side to the atmosphere through the inlet device, and on the other side to the inlet of the mixing chamber, an adjustable nozzle is installed at the outlet of the high-pressure channel of the gas ejector, a water manifold with nozzles directed along the air flow is located in the mixing chamber, a heat exchange device is located in the diffuser through which fuel is supplied to the main combustion chamber of the engine, the relative reduced speed of the rotor is determined by the formula Where - relative physical rotation frequency, - physical rotation frequency, - physical rotation frequency at cruise flight mode, - braking temperature of the air at the compressor inlet in the cruising flight mode, - braking temperature of air at the compressor inlet. 2. A hypersonic turbojet engine according to paragraph 1, characterized in that the pressure increase ratio in the compressor is no more than five. 3. A hypersonic turbojet engine according to claim 1, characterized in that the mixing chamber is cylindrical. 4. A hypersonic turbojet engine according to paragraph 1, characterized in that the turbocompressor working blades are single-crystal. 5. A hypersonic turbojet engine according to paragraph 1, characterized in that the water pressure in the water manifold is more than 10 MPa. 6. A hypersonic turbojet engine according to paragraph 1, characterized in that the engine turbine is single-stage. 7. A method for regulating a hypersonic turbojet engine comprising a supersonic inlet device, a turbocharger, an afterburner, an outlet device, in which a gas ejector with a mixing chamber and a diffuser is located between the inlet device and the compressor, the high-pressure channel of the gas ejector is connected on one side to a cavity located behind the turbine, and on the other side is connected to the mixing chamber, the low-pressure channel of the gas ejector is connected on one side to the atmosphere through the inlet device, and on the other side to the inlet of the mixing chamber, an adjustable nozzle is installed at the outlet of the high-pressure channel of the gas ejector, a water manifold with nozzles directed along the air flow is located in the mixing chamber, a heat exchange device is located in the diffuser through which fuel is supplied to the main combustion chamber of the engine, the relative reduced speed of the rotor is determined by the formula Where - relative physical rotation frequency, - physical rotation frequency, - physical rotation frequency at cruise flight mode, - braking temperature of the air at the compressor inlet in the cruising flight mode, - the braking temperature of the air at the compressor inlet, which consists of maintaining a constant relative reduced rotor speed. 8. A method for regulating a hypersonic turbojet engine according to paragraph 7, characterized in that at subsonic flight speeds up to M=0.9, the stagnation temperature of the air at the compressor inlet equal to the braking temperature of the outside air the relative physical rotation frequency is equal to relative reduced frequency 9. The method for regulating a hypersonic turbojet engine according to paragraph 7, characterized in that at flight speeds greater than M = 0.9 and less than the cruising speed, the stagnation temperature of the air at the compressor inlet is equal to the stagnation temperature of the air at the cruising flight mode - maintained by pressurizing the compressor with hot gas, the relative physical rotation frequency is equal to the relative reduced rotation frequency, upon reaching the cruising flight speed, the pressure difference in the turbine increases to the maximum, 10. A method for regulating a hypersonic turbojet engine according to paragraph 7, characterized in that at flight speeds exceeding the cruising speed, the physical rotation frequency increases in proportion to the square root of the air stagnation temperature at the compressor inlet until the gas temperature limit in front of the turbine is reached, after which water is supplied to the mixing chamber in an amount ensuring the maintenance of a constant air temperature at the outlet of the mixing chamber; when the water flow limit is reached, the afterburner is activated.