JPH07247908A - Coupled propulsion device possible to commonly use atmosphere or space - Google Patents

Abstract

PURPOSE:To provide a propellant engine capable of freely, simultaneously or selectively driving each propulsion mode out of a turbo, a ram and a rocket. CONSTITUTION:A variable nozzle 12 is build up between a taper forming part 27 and an outside case 4 by feely setting the axiial position of a thin-tip wide-rot nozzle 16 integrally formed with a combustion chamber fitted into a fuel injector 18 by a movable means 17, an annular combustion chamber 6 whose head an tail parts are provided with an air intake device 2 and an exhaust nozzle 13 respectively is formed by extending the taper forming part 27 to the outside case 4 so as to divide the outside case 4 from a center body 7, and a turbo pump 11 for supplying fuel is also provided. Hereby, each propulsion mode of a turbo, a ram and a rocket can be individually or simultaneously driven by retaining the singularity of a rocket combustion chamber so that the minimum single structural body for forming a stepless neck cross section in fuel gas for an atmospheric engine is realized by the most simple way of axial movement only.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a joint propulsion engine that can be commonly used in the atmosphere or the universe.

[0002]

2. Description of the Related Art In order to cope with various neck cross-sections for a combustion gas flow suitable for a difference in propulsion system such as turbo, ram, rocket, etc. and external operating conditions, dismantle and transfer each part of a tapered divergent nozzle, etc. Because of the complicated operation,
There is a problem with stability and reliability during ultra-high speed flight, and the entire tapered divergent nozzle is moved without disassembling or moving each part,
The method of changing the cross-section of the neck portion loses the unity of the rocket combustion chamber and obliges a complicated structure such as an annular injection port (Japanese Patent Laid-Open No. 1051053 / 52-5620).
9). This is inevitable because of this movement to adapt the combustion gas passage of the atmospheric engine to the directional quantity (vector) of the rocket combustion gas. In order to obtain an inexpensive, stable and reliable device, it is desirable to limit the forward / backward movement only on the shaft.

[0003]

The simplest on-axis transition of the smallest unitary structure to maintain the unity of the rocket combustion chamber and to provide a stepless neck cross-section for the combustion gases of the atmospheric engine. Providing a combined propulsion engine that can be realized only by itself and that each propulsion method of turbo, ram, rocket is reversible, multiple propulsion methods can be driven simultaneously or individually, and that the ram propulsion method can be used from takeoff start It is something to do.

[0004]

Therefore, the present invention is a combined propulsion engine that can commonly use the atmosphere or space,
Atmospheric engines that operate with atmospheric oxidants and liquids, or
The atmospheric engine comprises a rocket engine operated by a gaseous propellant, the atmospheric engine having an annular combustion chamber (6) defined by a central body (7) and an outer case (4). The rocket engine is equipped with an air intake device (2) for sending air, and an exhaust nozzle (13) for discharging combustion gas of the annular combustion chamber (6) is provided at an axial rear portion of the central body (7). Combustion chamber (10)
And a turbopump (11) for supplying a propellant to the rocket combustion chamber (10), in a joint propulsion engine commonly usable in the atmosphere or space, with its axis aligned with the central fuselage. The rocket engine combustion chamber (1
0) is integrally formed with the tapered divergent nozzle (16) of a constant length, and is fitted concentrically and substantially airtightly on the outer or inner circumference of the fixed fuel injector (18), and by the movable means (17), The axial position is adjustable, and the divergent trailing edge (21) of the tapered divergent nozzle (16) of the rocket engine and the neck peripheral edge (19) of the exhaust nozzle (13) of the combustion gas of the rocket engine A combustion chamber (10) of the rocket engine in order to ensure flow continuity, while adapting the flow cross section of the combustion gas from the annular combustion chamber (6) to different internal and external operating conditions of the engine; The divergent trailing edge portion (21) of the tapered divergent nozzle (16) integrally formed with the neck peripheral edge portion (1) of the exhaust nozzle (13).
The variable nozzle (12) that forms the optimum neck cross section for the combustion gas from the annular combustion chamber (6) is configured by making the distance from the annular combustion chamber (6) adjustable by the movable means (17). It is a solution.

[0005]

The essential advantage of the present invention is to obtain a stepless neck cross section by adjusting the mechanical position on the shaft and to freely adjust the pressure of combustion gas or non-combustion gas from the rocket combustion chamber. The combustion gas from the annular combustion chamber, which has passed through the neck section flowing into the exhaust nozzle, forms an inner wall in which the gas jet pressure from the rocket combustion chamber can fluctuate with respect to the outer case to which the exhaust nozzle is fixed. After passing through the variable nozzle that forms a predetermined neck cross section, the combustion gas from can be guided to various appropriate jet vectors with minimum turbulence, and the energy efficiency of the propelling gas can be enhanced.

[0006]

EXAMPLES Next, based on examples, referring to the drawings,
The present invention will be described in detail. FIG. 1 shows a combined propulsion engine that can be commonly used in the atmosphere or space, and this combined propulsion engine (1) has an air compressor (3) for compressing air supplied from an air intake device (2). Outer case (4)
And a central body (7) which is supplied with air compressed by an air compressor (3) and which defines an annular combustion chamber (6) with a fuel injector (5) with the outer case (4).

The air compressor (3) is driven by the turbine (8). This allows the choice to be made with atmospheric oxidant and with the gas generator (9). Figure 1
Supplies propellant (40) and oxidizer (28) to gas generator (9)
A turbo pump (11) for supplying the rocket combustion chamber (10) and the annular combustion chamber (6) and its flow path system are shown. The combustion gas generated in the annular combustion chamber (6) has a taper-shaped portion (27) with the outer case (4) extending rearward.
Is ejected to the exhaust nozzle (13) by being accelerated through the variable nozzle (12) formed by forming

The air compressor (3) and its related devices are omitted in the ramjet propulsion system, and the air compressor (3) keeps air at a compression ratio of 1 while freely rotating. This air is supplied to the annular combustion chamber (6) and also from the inlet (29) to the central gas pipe (15) when the flap opening / closing device (14) is open. The flap opening / closing device (14) is moved rearward by an operating pneumatic turbine (not shown) to move the moving part (14-1), the lattice plate (14-2), the closed door (14-3) and the closed door bar (14-). 4)
Consists of. The central gas pipe (15) can share the drive gas (30) for driving the turbine (8) with the annular combustion chamber (6) by adjusting the opening and closing of the flap opening and closing device (14). When the closing door (14-3) is closed, the whole amount is discharged from the lattice plate (14-2) into the annular combustion chamber (6), and when the moving part (14-1) is closed, the whole amount is the central pipe (15). Flows into the space between the shut-off bars (14-4). The quantitative distribution of these exhaust gases is based on the Suehiro trailing edge (2
Exhaust nozzle for forming a proper jet vector (31) after the combustion gas in the annular combustion chamber (6) whose discharge pressure from 1) is discharged to the exhaust nozzle (13) is discharged from the variable nozzle (12) It functions as an inner case (33) with respect to the outer case (32) of (13). Variable nozzle (1
In 2), a tapered divergent nozzle (16) integrally formed with a rocket combustion chamber (10) is formed by an actuator (17) to form a neck cross section adapted to different propulsion systems (turbo, ram) by a fuel injector (17). 18) The tapered divergent nozzle (16) formed integrally with the rocket combustion chamber and the neck cross-section peripheral edge end (19) of the exhaust nozzle (13) by moving the position on its axis rearward while being fitted in 18). The distance from the trailing edge (21) of the divergent end of the nozzle can be adjusted freely, and a variable nozzle (12) optimal for discharging combustion gas from the annular combustion chamber (6) is provided.

The fuel injector (18) includes a gas central tube (1
5) occupies its central position and has fuel and oxidant injection ports (20) on both sides thereof. Therefore, as an optimal one that can handle either liquid or gas as fuel, Japanese Patent Publication No. 1-1.
A 7000 "jet nozzle" (same as the applicant) is used. That is, "a high-pressure liquid jet port and a high-pressure gas jet port concentrically outside the high-pressure liquid jet port are provided so that their jetting directions are parallel to each other, and the high-pressure liquid jet port is retracted from the nozzle tip opening by a predetermined distance. And another small-diameter high-pressure gas injection port (center gas pipe 15-) between the rear of the central axis of the high-pressure liquid injection port and the injection port.
1) is provided and high-pressure gas is introduced ”.

The fuel injector (18) is fixed to the central body (7), and a tapered divergent nozzle (16) formed integrally with the rocket combustion chamber is fitted into the outer periphery of the fuel injector (18) in a substantially airtight manner. Then, if you move backward on the axis to the limit,
The divergent trailing edge (21) of the tapered divergent nozzle (16) integrally formed with the rocket combustion chamber at the neck peripheral edge (19) of the exhaust nozzle (13) continues the flow of the gas flowing therein. This joining completes the transition to rocket propulsion because the dimensions are guaranteed. In the rocket propulsion system, the expansion nozzle (22) is further held by the movable means (23) on the guide rail (24) and moved rearward, and the expansion nozzle (22) is moved to the divergent opening cross section (25) of the exhaust nozzle (13). The tapered cross section (26) is dimensioned for joining. In the rocket propulsion system, the flap opening / closing device (14) is closed and the variable nozzle (12) is closed, so that no backflow of rocket fuel gas occurs.

The annular combustion chamber (6) is made up of multiple parts to facilitate chamber strength and fabrication. The annular combustion chamber (6) is of circular cross section and is constituted by a plurality of individual chambers (6 ') arranged substantially adjacent to each other in a ring around the axis of the central body (7). (34) is its injection port, and its injection direction is parallel to the injection direction of the fuel injection port (20) of the fuel injector (18), and its opening position is projected from the fuel injection port (20). ing. In combination with the position of the high-pressure gas injection port (15-1) of the central gas pipe (15), the meaning of this is the configuration of another application of the Japanese Patent Publication No. 1-17000 "jet nozzle" mentioned earlier. Is shown. The divergent opening cross section of the exhaust nozzle (13) (2
5) corresponds to "nozzle tip opening".

The combined propulsion engine of the present invention can be adapted to the required continuous propulsion power, from a turbo propulsion system operating an atmospheric engine to a ram propulsion system, and
It is possible to operate the rocket system and the turbo system at the same time in order to obtain the maximum propulsion power while continuously changing to the rocket propulsion system. In this case, the gas from the rocket engine does not compensate for the pressure drop of the combustion gas (36) from the annular combustion chamber (6) that results from the rare air in the atmosphere, and the gas from the inner wall of the exhaust nozzle (13) Optimal propulsion efficiency is obtained by preventing separation and adjusting the variable nozzle (12) at the same time.

If the rocket engine is operated even after the transition to the ram system, the taper forming portion (27) causes an ejector action to further promote the intake of the atmosphere. This means that it is possible to operate the ram propulsion system from the start of takeoff. FIG. 2 shows a transition to a ram propulsion system in which the tapered divergent nozzle (16) brings its axial position closer to the neck peripheral edge (19) of the exhaust nozzle (13),
A reduction in the throat cross section of the variable nozzle (12) is obtained.

FIG. 3 shows a state of transition to the rocket propulsion system. The divergent trailing edge portion (21) of the tapered divergent nozzle (16) formed integrally with the rocket combustion chamber is the exhaust nozzle (1
3) is joined to the peripheral edge portion (19) of the neck portion, and the expansion nozzle (22) is moved rearward so that the tapered cross section (26) is joined to the divergent opening cross section (25) of the exhaust nozzle (13). There is.

[0015]

[Effect of the Invention] The minimum unitary structure for maintaining the unity of the rocket combustion chamber and constructing a stepless neck cross-section for the combustion gas of the atmospheric engine is realized by the simplest axial transition. At the same time, the turbo, ram, and rocket propulsion methods are reversible, and multiple propulsion methods can be driven simultaneously or individually, and the ram propulsion method can be used from the time of takeoff and launch.

[Brief description of drawings]

FIG. 1 is an axial cross-sectional view showing one example of a coupled propulsion engine of the present invention, showing that it is in a low altitude flight state and also uses a rocket engine in some cases.

FIG. 2 is a sectional view showing a state in which the throat section of the variable nozzle (12) is smaller than that in FIG.

FIG. 3 is a cross-sectional view of a state in which the transition to the rocket propulsion system is completed.

[Explanation of symbols]

 2 Air Intake Device 3 Air Compressor 4 Outer Case 5 Fuel Injector 6 Annular Combustion Chamber 7 Central Fuselage 10 Rocket Combustion Chamber 11 Turbo Pump 12 Variable Nozzle 13 Exhaust Nozzle 16 Tapered Wide-end Nozzle 17 Movable Means 18 Fuel Injector 19 Exhaust nozzle neck peripheral edge

Claims (1)

[Claims] 1. A coupled propulsion engine that can be commonly used in the atmosphere or the universe, comprising an atmospheric engine that operates with an oxidizing agent in the atmosphere and a rocket engine that operates with a liquid or gaseous propellant, Atmospheric engine, central fuselage (7)
And an outer case (4) to define an annular combustion chamber (6) and an air intake device (2) for sending air to the annular combustion chamber (6). An exhaust nozzle (13) for discharging combustion gas is provided at an axial rear part of the central body (7), and the rocket engine includes a rocket combustion chamber (10) and a propellant for the rocket combustion chamber (1).
In a joint propulsion engine which can be commonly used in the atmosphere or the universe, and which is equipped with a turbo pump (11) for supplying 0), a rocket engine arranged with its axis aligned with the axis of the central fuselage The combustion chamber (10) is integrally formed with a tapered divergent nozzle (16) having a fixed length, and is fitted concentrically and substantially airtightly on the outer or inner periphery of a fixed fuel injector (18), and a movable means (17). ) Allows the axial position to be adjusted, and the tapered divergent nozzle of the rocket engine (16)
The suehiro trailing edge (21) and the neck peripheral edge (19) of the exhaust nozzle (13) are dimensioned to ensure continuity of the flow of combustion gas of the rocket engine, while the annular combustion box (6)
A divergent trailing edge of a tapered divergent nozzle (16) formed integrally with the combustion chamber (10) of the rocket engine to adapt the flow cross section of combustion gas from the engine to different internal and external operating conditions of the engine ( 21) and the exhaust nozzle (1
The variable nozzle (3) that forms an optimum neck cross section for the combustion gas from the annular combustion chamber (6) by making the distance between the neck peripheral edge (19) and the movable means (17) adjustable. 12) A combined propulsion engine, characterized in that