CN105736178B - Combined cycle engine - Google Patents

Abstract

The invention provides a combined cycle engine, which includes: a supersonic air intake mechanism, having a first gas inlet open in the axial direction and communicating with the outside; a continuously rotating detonation engine, fixedly connected to the downstream of the supersonic air intake mechanism ; and a turbojet engine fixedly connected downstream of the supersonic air intake mechanism. Wherein, both the continuously rotating detonation engine and the turbojet engine are in controlled communication with the supersonic air intake mechanism to controlly introduce external air through the first gas inlet to work. The combined cycle engine of the present invention can complete the mode conversion within the range of the working envelope, provide stable thrust for the aircraft, and then enable the aircraft to achieve hypersonic flight above Mach 5.0 in the atmosphere and to take off and land horizontally. In addition, based on the continuous rotation detonation technology, the specific impulse of the engine is increased, the thermal efficiency of combustion is improved, and the economy and reliability of the engine are improved.

Description

combined cycle engine

technical field

The invention relates to the technical field of aerospace power, in particular to a combined cycle engine.

Background technique

Near-space hypersonic vehicles with a maximum flight speed of around Mach 5.0 have advantages such as superior high-altitude and high-speed characteristics, flexible maneuverability, and high penetration probability, and have very important potential military value. In addition, with the advancement of missile defense technology, the penetration and survivability of traditional missiles are also greatly threatened, and there is an urgent need to improve mobility through high-speed propulsion devices. However, this type of aircraft has the characteristics of wide working envelope and complex and changeable flight conditions. It is difficult for a single type of propulsion device to complete all the work requirements independently. Therefore, combined power technology emerged as the times require and quickly attracted people's attention. The two types of combined cycle engines that are currently being studied are the rocket ramjet combined cycle engine and the turbo ramjet combined cycle engine. The former is not economical because it does not effectively use the oxygen in the atmosphere, and it is currently difficult to meet the requirements of repeated use. For the latter, The biggest problem is the mode conversion. The maximum working Mach number of the turbine engine is lower than the effective starting Mach number of the ramjet engine, so that when the combined engine switches from the turbine mode to the ramjet mode, the thrust cannot meet the requirements.

Contents of the invention

In view of the problems existing in the background technology, an object of the present invention is to provide a combined cycle engine, which can successfully complete the mode conversion within the working envelope range, provide stable thrust for the aircraft, and then enable the aircraft to achieve Mach 5.0 in the atmosphere More than hypersonic flight and can take off and land horizontally.

Another object of the present invention is to provide a combined cycle engine, which increases the specific impulse of the engine, solves the problem of unstable combustion, further improves the thermal efficiency of combustion, and improves the economy and reliability of the engine.

In order to achieve the above object, the present invention provides a combined cycle engine, which includes: a supersonic air intake mechanism, having a first gas inlet communicated with the outside along the axial direction; a continuously rotating detonation engine, fixedly connected to the supersonic downstream of the air intake mechanism; and a turbojet engine fixedly connected downstream of the supersonic air intake mechanism. Wherein, both the continuously rotating detonation engine and the turbojet engine are in controlled communication with the supersonic air intake mechanism to controlly introduce external air through the first gas inlet to work.

The beneficial effects of the present invention are as follows:

In the combined cycle engine according to the present invention, both the continuously rotating detonation engine and the turbojet engine are in controlled communication with the supersonic air intake mechanism, so that the mode conversion can be successfully completed within the working envelope of the combined cycle engine, and the aircraft Provide stable thrust, and then enable the aircraft to achieve hypersonic flight above Mach 5.0 in the atmosphere and be able to take off and land horizontally. In addition, based on the continuous rotation detonation engine, the continuous rotation detonation technology is adopted, which improves the specific impulse of the engine, solves the problem of combustion instability, improves the thermal efficiency of combustion, and improves the economy and reliability of the engine.

Description of drawings

Fig. 1 is a working schematic diagram of a working mode of the combined cycle engine according to the present invention, wherein the second gas inlet and the third gas inlet are all partially opened, and the window is in a closed state, and the arrow indicates the air flow direction;

Fig. 2 is a schematic diagram of the circumferential position of the fuel nozzle of the continuously rotating detonation engine in Fig. 1;

Fig. 3 is a schematic diagram of the circumferential position of the igniter of the continuously rotating detonation engine in Fig. 1;

Fig. 4 is a working schematic diagram of another working mode of the combined cycle engine according to the present invention, wherein the second gas inlet is closed, the third gas inlet is opened and the window is in the state of inwardly bleed air, and the arrow indicates the airflow direction;

5 is a working schematic diagram of another working mode of the combined cycle engine according to the present invention, wherein the second gas inlet is opened, the third gas inlet is closed and the window is in a closed state, and the arrow indicates the air flow direction;

Fig. 6 is a working schematic diagram of the combined cycle engine in the same working mode as Fig. 1, wherein both the second gas inlet and the third gas inlet are partially opened, and the window is in an outward exhaust state, and the arrow indicates the airflow direction.

Wherein, the reference signs are explained as follows:

1 supersonic intake mechanism 25 igniter

11 First gas inlet 26 Fuel nozzle

12 first casing 3 turbojet engine

13 central body 31 third shell

131 Precursor 311 Head

132 rear body 312 tail

1321 Increased diameter arc surface 32 The third gas inlet

1322 Reduced diameter arc surface 33 The third gas outlet

1323 reduced diameter truncated cone face 4 first adjustment mechanism

14 intake runner 41 valve

15 windows 42 cylinders

2 continuous rotation detonation engine 421 cylinder

21 Second housing 422 Piston rod

22 annular combustion chamber 5 second adjustment mechanism

23 Second gas inlet 51 Vane

24 Second gas outlet

Detailed ways

The combined cycle engine according to the present invention will be described in detail below with reference to the accompanying drawings.

With reference to Fig. 1 to Fig. 6, the combined cycle engine according to the present invention comprises: supersonic air intake mechanism 1, has the first gas inlet 11 that is opened along the axial direction that communicates with the outside; The downstream of the sonic air intake mechanism 1; and the turbojet engine 3, which is fixedly connected to the downstream of the supersonic air intake mechanism 1. Wherein, both the continuously rotating detonation engine 2 and the turbojet engine 3 are in controlled communication with the supersonic air intake mechanism 1 so as to controlly introduce external air through the first gas inlet 11 to work.

In the combined cycle engine according to the present invention, both the continuously rotating detonation engine 2 and the turbojet engine 3 are in controlled communication with the supersonic intake mechanism 1, so that the mode conversion can be successfully completed within the working envelope of the combined cycle engine , to provide stable thrust for the aircraft, so that the aircraft can achieve hypersonic flight above Mach 5.0 in the atmosphere and can take off and land horizontally. In addition, based on the continuous rotation detonation engine 2, the continuous rotation detonation technology is adopted, which improves the specific impulse of the engine, solves the problem of combustion instability, and thus improves the thermal efficiency of combustion, and improves the economy and reliability of the engine.

According to the combined cycle engine of the present invention, in one embodiment, referring to FIG. 1 and FIG. 4 to FIG. 6 , the supersonic air intake mechanism 1 may include: a first casing 12 and a central body 13 . The central body 13 may include: a front body 131 protruding from the first housing 12; The first gas inlet 11 is formed between the outer wall surface of the rear body 132 and the first casing 12 , and an intake flow channel 14 communicating with the first gas inlet 11 is formed. It is supplemented here that since the first gas inlet 11 is narrower than the intake flow passage 14, the intake flow passage 14 formed between the outer wall surface of the rear body 132 and the first casing 12 is a compression passage, especially in super Under the condition of sonic flight, the incoming air can form a shock wave train in the air intake channel 14 to decelerate and boost the pressure.

In one embodiment, referring to FIG. 1 and FIGS. 4 to 6 , the front body 131 is tapered. The rear body 132 may have: an increasing diameter arc surface 1321, the diameter gradually increases from the axial downstream end of the front body 131 toward the axial downstream direction; a decreasing diameter arc surface 1322, from the axial downstream end of the increasing diameter arc portion to the axial downstream and the reduced diameter truncated cone surface 1323 , whose diameter gradually decreases from the axially downstream end of the reduced diameter arc portion to the axially downstream direction until the turbojet 3 .

In an embodiment, referring to FIG. 1 to FIG. 6 , the continuous rotation detonation engine 2 may include: a second casing 21 fixedly connected to the axially downstream end of the first casing 12 .

In one embodiment, referring to FIGS. 1 to 6 , the turbojet engine 3 may comprise: a third housing 31 fixedly connected to the axially downstream end of the rear body 132 of the central body 13 of the supersonic air intake mechanism 1 . Wherein, the head 311 of the third housing 31 protrudes into the second housing 21 to form a continuous rotation detonation between the outer wall surface of the head 311 of the third housing 31 and the inner wall surface of the second housing 21 Annular combustion chamber 22 of engine 2 . The rotary detonation combustion based on the fuel in the annular combustion chamber 22 is equal-volume combustion, thereby significantly improving thermal efficiency, thereby reducing the fuel consumption rate of the continuous rotary detonation engine 2 and improving fuel economy.

In one embodiment, referring to Fig. 1 and Fig. 4 to Fig. 6, the continuous rotation detonation engine 2 may have: a second gas inlet 23; Exhaust gases from internal combustion. The turbojet engine 3 may have: a third gas inlet 32; and a third gas outlet 33 arranged at the tail portion 312 of the third housing 31 of the turbojet engine 3 and communicating with the interior of the turbojet engine 3 to discharge the turbojet engine 3 Gases inside. The combined cycle engine can also include: a first regulating mechanism 4, which controls the opening or closing of the second gas inlet 23 so that the annular combustion chamber 22 of the continuous rotation detonation engine 2 communicates with the intake air passage 14 of the supersonic air intake mechanism 1 or Disconnect; and the second regulating mechanism 5, which controls the opening or closing of the third gas inlet 32 so that the inside of the turbojet engine 3 is communicated with or disconnected from the intake air passage 14 of the supersonic air intake mechanism 1.

It is supplemented here that when the incoming flow velocity is less than Mach 1.5, the incoming air enters the intake air channel 14 through the first gas inlet 11 to be compressed, and the temperature and pressure of the compressed air are not very high, which is not conducive to The formation of the detonation wave can be achieved by controlling the first regulating mechanism 4 to close the second gas inlet 23, so that the continuous rotation detonation engine 2 does not work, and only the turbojet engine 3 works, as shown in FIG. 4 . At this time, the opening degree of the third gas inlet 32 can be increased by controlling the second regulating mechanism 5 , so that all incoming air enters the turbojet engine 3 .

When the incoming flow velocity reached Mach 1.5, the continuous rotation detonation engine 2 started to work, and gradually enlarged the opening degree of the second gas inlet 23 by controlling the first regulating mechanism 4; meanwhile, the performance of the turbojet engine 3 had been attenuated relatively. The opening degree of the third gas inlet 32 can be gradually reduced by controlling the second adjusting mechanism 5, as shown in FIG. 1 . Under this incoming flow speed state, two kinds of motors work simultaneously.

When the incoming flow velocity was equal to Mach 2.5, the third gas inlet 32 was completely closed, and the turbojet engine 3 completely stopped working. And the second gas inlet 23 is opened to the maximum, as shown in Figure 5, all the incoming air enters in the annular combustion chamber 22 of the continuous rotary detonation engine 2 through the second gas inlet 23, and forms a mixture with fuel, through the hereinafter described After the above-mentioned igniter 25 is ignited, a detonation wave is formed in the annular combustion chamber 22, and then the exhaust gas is discharged through the second gas outlet 24, thereby generating thrust.

In an embodiment, referring to FIG. 1 and FIG. 4 to FIG. 6 , the first adjustment mechanism 4 may include: a valve 41 and a cylinder 42 . The valve 41 is arranged in the second casing 21 of the continuously rotating detonation engine 2 and is slidably mounted on the front of the third casing 31 of the turbojet engine 3 in a position close to the first casing 12 . The cylinder 42 includes: a cylinder body 421, one end of which is pivotally connected to the portion behind the axial direction of the valve 41 of the second housing 21; Connect to valve 41.

In one embodiment, referring to FIG. 1 and FIG. 4 to FIG. 6 , the second adjustment mechanism 5 may include: a plurality of vanes 51 , and the rear body 132 near the central body 13 is radially installed in the front portion of the third housing 31 . Wherein, each blade 51 can rotate around its own installation shaft (not shown). When each blade 51 rotates around its respective installation axis and the adjacent blades 51 are separated from each other, the third gas inlet 32 is opened; when each blade 51 rotates around its respective installation axis and the adjacent blades 51 are closely connected to each other, The third gas inlet 32 is closed.

In an embodiment, referring to Fig. 1 and Fig. 4 to Fig. 6, the supersonic air intake mechanism 1 may further include: a window 15, which is arranged through the first housing 12, and communicates with the air intake channel 14 in a controlled manner, so as to The air intake channel 14 inputs supplementary external air or discharges the gas in the air intake channel 14 to the outside. It is supplemented here that when the incoming flow velocity is less than Mach 0.6, since the first gas inlet 11 is narrow and the incoming flow is subsonic, the incoming air is not enough to meet the working requirements of the engine. At this time, the supersonic air intake mechanism 1 The window 15 of the window 15 can be adjusted to the state of bleed air, so that the outside air can be introduced into the air intake channel 14 through the window 15 to supplement the air flow, as shown in FIG. 4 . And when intake distortion occurs, the window 15 can be adjusted to a deflated state to lead the air in the intake runner 14 out to the outside, thereby expanding the stable operating range of the engine, as shown in FIG. 6 .

In one embodiment, the window 15 is several rounds of grooves 151 opened at the rear of the first housing 12 in the circumferential direction. The side of the channel 151 connected to the atmosphere is provided with an adjusting piece (not shown). The adjusting piece can rotate along its respective installation axis, thereby realizing the control of the flow rate and direction of the airflow. Atmospheric differential control.

In one embodiment, referring to FIG. 1 and FIG. 3 to FIG. 6 , the continuous rotation detonation engine 2 may further include: an igniter 25 disposed in the annular combustion chamber 22 of the continuous rotation detonation engine 2 for controlling the incoming The gas (that is, the mixture of fuel and air) in the annular combustion chamber 22 is ignited, and then the gas detonates and combusts.

In one embodiment, referring to Fig. 1, Fig. 2 and Fig. 3 to Fig. 6, the continuous rotation detonation engine 2 may also have: a plurality of fuel nozzles 26, which penetrate the second housing 21 in the radial direction and the annular combustion chamber 22 Corresponding part, and distributed along the circumferential direction of the second housing 21 , each fuel nozzle 26 communicates with the inside of the turbojet engine 3 and the annular combustion chamber 22 , so that fuel enters the annular combustion chamber 22 through each fuel nozzle 26 . It is supplemented here that a plurality of fuel nozzles 26 are distributed along the circumferential direction of the second housing 21, so that the direction in which the fuel enters the annular combustion chamber 22 and the flow direction of the air (as an oxidant) that enters through the second gas inlet 23 can be made It is 90°, which helps the fuel and air to fully mix.

In an embodiment, the combined cycle engine may further include: a control system (not shown), communicating with the cylinder 42 of the first regulating mechanism 4 , the second regulating mechanism 5 and the window 15 .

In one embodiment, the combined cycle engine can be used as the power system of a hypersonic vehicle in near space, the power system of the first stage of a two-stage launch system, or the power system of a hypersonic civil aircraft and an aerospace aircraft.

Claims (7)

1. a kind of combined cycle engine, which is characterized in that including：

Supersonic speed admission gear (1) has the first gas inlet (11) open vertically connected with outside；

Continuous rotation detonation engine (2) is fixedly connected on the downstream of supersonic speed admission gear (1)；And

Turbojet (3) is fixedly connected on the downstream of supersonic speed admission gear (1)；

Wherein, continuously rotate detonation engine (2) and turbojet (3) with supersonic speed admission gear (1) controlled company It passes to and works via the controlled introducing extraneous gas of first gas inlet (11)；

Supersonic speed admission gear (1) includes：

First shell (12)；And

Centerbody (13), including：

Precursor (131) stretches out in first shell (12)；

Body (132) afterwards are contained in first shell (12), boundary position and first shell of the rear body (132) with precursor (131) (12) first gas inlet (11) is formed,

And it is formed with the air inlet connected with first gas inlet (11) between the outside wall surface of rear body (132) and first shell (12) Runner (14)；

Supersonic speed admission gear (1) further includes：Window (15), perforation is arranged in first shell (12), with air inlet runner (14) Controlled connection, to be discharged to the extraneous gas of air inlet runner (14) input supplement or by the gas in air inlet runner (14)；

Window (15) is a few conduits (151) circumferentially outputed at first shell (12) rear portion, conduit (151) and big gas phase One side even is provided with adjustment sheet, and adjustment sheet rotate along respective installation axle, the angle of adjustment sheet by air inlet runner (14) with outside Boundary's atm difference control.

2. combined cycle engine according to claim 1, which is characterized in that continuously rotating detonation engine (2) includes：

Second shell (21) is fixedly connected on the axial downstream end of first shell (12).

3. combined cycle engine according to claim 2, which is characterized in that turbojet (3) includes：

3rd housing (31) is fixedly connected on the axial downstream of the rear body (132) of the centerbody (13) of supersonic speed admission gear (1) End；

Wherein, the head (311) of the 3rd housing (31) is stretched into second shell (21), on the head of the 3rd housing (31) (311) toroidal combustion chamber of continuous rotation detonation engine (2) is formed between outside wall surface and the internal face of second shell (21) (22)。

4. combined cycle engine according to claim 3, which is characterized in that

Continuous rotation detonation engine (2) has：

Second gas inlet (23)；And

Second gas exports (24), connects to discharge the exhaust gas in toroidal combustion chamber (22) after burning with toroidal combustion chamber (22)；

Turbojet (3) has：

Third gas entrance (32)；And

Third gas exports (33), is arranged at the afterbody (312) of the 3rd housing (31) of turbojet (3), and and whirlpool The inside of wheel jet engine (3) connects to discharge the internal gas of turbojet (3)；

Combined cycle engine further includes：

First adjusting mechanism (4), control second gas inlet (23) are opened or closed so as to continuously rotate detonation engine (2) Toroidal combustion chamber (22) connect or disconnect with the air inlet runner (14) of supersonic speed admission gear (1)；And

Second adjusting mechanism (5), control third gas entrance (32) are opened or closed so that turbojet (3) is internal With the connection or disconnection of the air inlet runner (14) of supersonic speed admission gear (1).

5. combined cycle engine according to claim 4, which is characterized in that the first adjusting mechanism (4), including：

Valve (41) is arranged in the second shell (21) of continuous rotation detonation engine (2), and is slidably mounted on turbojet The position of close first shell (12) on the forepart of 3rd housing (31) of engine (3)；And

Cylinder (42), including：

Cylinder body (421), one end are pivotally connected to the part of the axial rearward direction in valve (41) of second shell (21)；And

Cylinder body (421) is stretched into piston rod (422), one end, and the other end stretches out cylinder body (421) and is pivotally connected to valve (41).

6. combined cycle engine according to claim 4, which is characterized in that the second adjusting mechanism (5) includes：

Multiple blades (51), rear body (132) radially installed close to centerbody (13) is in the forepart of the 3rd housing (31)；

Wherein, each blade (51) can rotate around respective installation axle, when each blade (51) around respective installation axle rotate and When being separated from each other between adjacent blades (51), third gas entrance (32) is opened；When each blade (51) turns around respective installation axle When dynamic and adjacent blades (51) are closely joined together, third gas entrance (32) is closed.

7. combined cycle engine according to claim 1, which is characterized in that continuously rotate detonation engine (2) and also wrap It includes：

Igniter (25) is arranged in the toroidal combustion chamber (22) of continuous rotation detonation engine (2), for entering annular Combustion gas in combustion chamber (22) is lighted a fire, and then combustion gas detonating combustion；And

Multiple fuel nozzle ports (26) radially penetrate through the part corresponding with toroidal combustion chamber (22) of second shell (21), and edge Circumferentially distributed, each fuel nozzle ports (26) connection turbojet (3) inside and the toroidal combustion chamber of second shell (21) (22), so that fuel enters toroidal combustion chamber (22) via each fuel nozzle ports (26).