RU2623134C1 - Solid fueled integrated straight-jet engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: engine comprises an air-intake device (AID) with supply ductworks and air blow-off ducts, a compression chamber (CC) with housed within the front-body section channel grain of the sustainer propulsion stage solid-fuel, a flame holder, a sustainer propulsion stage solid-fuel grain and a cruise nozzle. The supply ductworks and air blow-off ducts link AID and CC at the front and back ends of the sustainer propulsion stage solid-fuel grain. The launching stage gain is implemented within the cylindrical case and is placed centrally with the circular clearance at the CC front end. An extra solid fuel ring charge is mounted on the launching stage outside housing face.

EFFECT: invention improves the density impulse of the engine unit launching stage and sustainer propulsion stage.

4 cl, 1 dwg

Description

The invention relates to jet technology, namely to combined ramjet engines, and can be used as a propulsion system of aircraft (LA).

One of the main directions in solving the problem of increasing the range and flight speed of aircraft with an intra-atmospheric operating zone is the use of integral ramjet engines (ramjet). Thanks to integration, the advantages of a solid propellant rocket engine (solid propellant rocket engine) in the launch stage and the high economic performance of ramjet engines in the march stage can be best realized in these engines.

As the starting stage, the solid propellant rocket engine is used, which accelerates the aircraft to the Mach number (M) of the inclusion of the march ramjet engine. There are a number of products for which it is required to ensure the minimum mass of discharged structural elements when switching from the acceleration mode of an aircraft to the operation mode of a main engine. The placement of the launch stage requires the allocation of significant volumes under strict dimensional constraints, therefore, modern layouts of propulsion systems are based on the maximum integration of the launch and flight stages.

Of interest for use as a marching engine in a combined ramjet engine is a ramjet engine for solid fuel (ramjet). In such an engine, the cruise mode of operation is realized when one or more charges of solid fuel are burned in the air stream. The advantages of ramjet engine include the relative simplicity of the design, high energy characteristics of solid fuels, increased reliability and the ability to automatically adjust within certain limits when changing flight performance.

Known propulsion system (US patent No. 4841724 IPC F02K 7/18), in which as the starting stage is used solid propellant rocket motors in a separate housing. The launch stage is connected to the march stage in the form of ramjet with the help of pyro-bolts. Upon reaching the required speed of acceleration and fuel production, the starting stage is separated from the propulsion system by the explosion of pyro-bolts, and the sustainer stage is launched.

The disadvantage of this solution is the large mass of discharged structural elements during the transition from the acceleration mode to the operation mode of the ram engine, which is not always permissible.

Known combined rocket and ramjet engine (US patent No. 5537815, IPC F02K 7/18), containing a gas generator with a charge of solid fuel, a four-pipe air intake and a combustion chamber, in which is located the starting non-nozzle rocket engine of solid fuel (BSRDT). BSRDT provides acceleration of the aircraft to the calculated value of the number M of the beginning of the operation of the main engine. After that, the combustion chamber serves to mix the gas generation products with the incoming air stream and their afterburning. This design ensures the absence of discharged in flight engine elements.

The disadvantage of this design is the large overall dimensions associated with the sequential placement of the starting and marching stages. In addition, the parameters of the gas generator do not depend on external conditions, and therefore there is an acute problem of regulating the flow of solid fuel gas generation products, which leads to a complication of the design.

Known propulsion system (L.S. Yanovsky, V.N. Alexandrov, etc. Integrated ramjet engines for solid fuels. M.: IKC "Akademkniga", 2006, S. 193, Fig. 4.2), in which the start -acceleration solid propellant rocket engine is built into the ramjet combustion chamber and remains in it until the complete completion of the propulsion system. This design provides a relatively small overall dimensions of the propulsion system.

The disadvantage of this design is the use of liquid fuel for the march stage, which requires a system for supplying and regulating the flow of liquid fuel. This leads to a decrease in its reliability and an increase in mass.

The closest analogue selected for the prototype is a combined solid-fuel ramjet (SFIRR, USA) for a long-range aviation anti-ship missile (VV Gavrilov, L.A. Klimenko and others. Foreign aircraft engines. M: Publishing House. Aviamir, 2005, p. 347, Fig. 1) with air bypass with built-in solid propellant rocket motor as a starting stage. A propulsion system comprises a combustion chamber with a channel charge of a solid propellant sustainer stage placed in it, a charge of solid fuel of the launch stage, a flame stabilizer, an air intake device with air supply and bypass channels, plugs for air intakes (not shown) and bypass channels, a mixing device, a launch and march nozzles.

The engine is equipped with an air intake device (VZU) with an arm-mounted segment-ring inlet and has a short curved channel for supplying air to the front of the combustion chamber, as well as an extended transition channel for bypassing part of the air flow through the openings in the combustion chamber located behind the end of the sustainer charge of hydrocarbon solid fuel. Air is introduced at an angle of 90 ° to the axis of the combustion chamber on a 90 ° segment around the circumference. Air for bypass is taken from the peripheral width sections of the annular segment inlet. To increase the strength and rigidity of the structure at the inlet portion of the air intake there are pylons connecting the lower and upper walls. The launch stage is integrated with the ramjet path and is a solidly bonded solid fuel charge located in the rear of the combustion chamber and in four grooves of the main fuel, and a launch nozzle. The placement of the charge in the grooves of the sustainer fuel provides good volumetric filling of the engine, improves the ignition of the sustainer fuel, and the ledges allow you to develop the initial combustion surface of the sustainer fuel. In acceleration mode, the air intake is closed by durable resettable plugs (not shown). To stabilize the combustion process and evenly distribute the air flow in the direct-flow path, an injector is located in front of the front end of the march fuel, and for good mixing of the primary combustion products of the charge with air, there is a mixing diaphragm behind the rear end of the march fuel charge, which ensures bypass air is mixed with the primary combustion products fuel in the first circuit of the combustion chamber.

The charge of solid fuel of the launch stage ensures the acceleration of the aircraft to the number M of the beginning of the operation of the main engine. In the marching mode, the air coming from the VZU is divided into two parts: the smaller one enters the solid fuel charge channel, and the rest is fed into the combustion chamber through the air bypass channels. The air entering the charge channel of the solid propellant sustainer stage is used for primary gasification of fuel. The resulting primary gasification products, enriched with combustible components, then enter the free volume of the combustion chamber, where they are mixed with air from the bypass channels and burned.

The technical solution with the placement of the solid fuel charge of the starting stage in the solid fuel charge channel allows to reduce the overall dimensions of the propulsion system, but it has several disadvantages.

The direct casting of solid fuel of the starting stage into the charge channel of the solid combustible sustainer stage significantly complicates the design.

The operation of the propulsion system in marching mode is possible only after the charge of the starting stage is completely burned out, which leads to some failure of thrust at the end of the acceleration mode and the beginning of the marching mode, which can lead to a decrease in the speed of the aircraft and non-launch of the VZU marching stage. To prevent such work, the mass of charge of solid fuel at the starting stage is assumed to be excessive.

In the marching mode, the fuel charge burns out non-uniformly along the length; at the exit from the combustion channel, the entrainment of solid fuel is greater than at the inlet.

A common combustion chamber for acceleration and marching conditions requires increased strength for operation at high pressure, which leads to a noticeable increase in the mass of the propulsion system. To operate the starting stage at low pressure, fuels with low energy consumption are used. In this case, the specific impulse is significantly reduced.

The invention is based on the following tasks:

- reduction in the mass of discharged structural elements during the transition from the acceleration mode to the operation mode of the march stage;

- improving the overall dimensions of the propulsion system;

- improving the reliability of the propulsion system;

- ensuring a smooth transition to marching mode;

- boosting the thrust of the propulsion system at the beginning of the march stage;

- ensuring uniform combustion of the charge of solid fuel on the marching mode.

The technical result of the claimed invention is to increase the specific impulses of the starting and marching stages of the propulsion system.

The tasks are solved in that the integrated ramjet engine on solid fuel (IPVRDT) contains an air intake device with air supply and bypass channels, a combustion chamber with a channel charge of a solid marching sustainer stage located in front of it, a flame stabilizer, a charge of solid fuel of the starting stage and marching nozzle. Moreover, the air supply and bypass channels connect the air intake device to the combustion chamber in the front and rear parts of the charge of the solid fuel march stage.

New in the invention is that the charge of solid fuel of the starting stage is made in a cylindrical body and placed centrally with an annular gap in front of the combustion chamber, and an additional ring charge of solid fuel is installed on the outer surface of the housing of the starting stage.

With this design, IPVRDT:

- the placement of the charge of solid fuel of the launch stage in a separate building centrally reduces the mass of discharged structural elements when switching from the acceleration mode of the aircraft to the operation mode of the sustainer stage due to the absence of a discharged nozzle;

- placement of the charge of solid fuel of the starting stage, the operating mode of which is carried out at a pressure of up to 10 MPa, in a separate housing centrally provides an improvement in the weight and size characteristics of the propulsion system by reducing the thickness of the walls of the combustion chamber of the march stage to the value necessary to ensure strength at operating pressure to 0, 7 MPa in march mode;

- placement of the charge of solid fuel of the launch stage in a separate housing centrally provides an increase in the specific impulse of the launch stage due to the allowable increase in operating pressure in the start-acceleration mode with an acceptable increase in the mass of the propulsion system;

- the coaxial arrangement of the launch stage housing and the charge of the solid propellant sustainer stage ensures reliable launch of the air intake device and the smooth transition of the aircraft in speed to the sustainer mode due to the possibility of launching the sustainer stage until the launch stage charge is completely burned out;

- burning an additional ring charge of solid fuel on the outer surface of the housing of the starting stage provides forcing the engine at the beginning of the sustainer mode of engine operation by redistributing the air flow between the engine circuits and increasing traction;

- the presence of an additional ring charge of solid fuel on the outer surface of the housing of the launch stage provides uniform combustion of the charge of solid fuel marching stage due to the additional heat flow to the surface of the charge in the front.

The development and refinement of the set of essential features of the invention for particular cases of its implementation is given below:

- the implementation of the outer surface of the additional charge with a taper angle of up to 30 degrees to the axis of the engine provides uniform burnout of the additional and main charges of the fuel marching stage. With an increase in the taper angle of more than 30 degrees, the flow path becomes excessively cluttered, which leads to a decrease in the recovery coefficient of the total pressure and, consequently, the engine thrust;

- the shift of the additional fuel charge relative to the front end of the charge of the sustainer stage along the axis towards the sustainer nozzle by a distance of 1 to 3 outer diameters of the housing of the starting engine provides the location of the additional charge in the region of high gas current density, which reduces the required mass of the additional fuel charge. When the additional charge is located near the end of the charge of the sustainer stage at a distance of less than one diameter of the starting engine housing, the input section of the flow path is cluttered. When the additional charge of fuel is displaced relative to the front end of the charge of the sustainer stage by a distance of more than three outer diameters in the front of the charge of the sustainer stage, high values of the gas current density and heat flux to the combustion surface are not provided;

- the manufacture of an additional charge of solid fuel with the addition of at least 10% metal particles (aluminum, magnesium, etc.) in the form of a fine powder provides a predetermined thrust boost at the beginning of the operation of the march stage by increasing the energy consumption of the fuel. The lower content of metal particles in the composition of the solid fuel does not provide the required increase in energy consumption of the fuel.

Thus, the objectives of the invention are solved:

- provided a reduction in the mass of discharged structural elements during the transition from the acceleration mode to the operation mode of the march stage;

- improved overall dimensions of the propulsion system;

- increased reliability of the propulsion system;

- provided the ability to force traction at the beginning of the march stage;

- Ensured a smooth transition to marching mode;

- ensured uniform combustion of the charge of solid fuel on the marching mode.

The present invention is illustrated by the following detailed description of the design and operation of an integral ramjet solid-fuel jet engine with reference to the drawing.

The solid-fuel integrated ramjet engine with solid fuel contains (see drawing) an air intake device 1 with channels 2, 3, respectively, for supplying and passing air, a combustion chamber 4 with a channel charge 5 of solid marching stage located in front of it, a flame stabilizer 6, a charge 7 solid fuel of the starting stage and the march nozzle 8. Moreover, the channels 2, 3 of supply and bypass of air connect the air intake device 1 to the combustion chamber 4 in the front and rear parts of the charge 5 of the march stage. The charge 7 of solid fuel starting stage is made in a cylindrical body 9 and is placed centrally with an annular gap in front of the combustion chamber 4. On the outer surface of the housing 9 of the starting stage an additional ring charge 10 of solid fuel is installed. The outer surface of the additional charge of solid fuel 10 has a taper angle of up to 30 degrees to the axis of the integral engine toward the nozzle 8 or air intake device 1. The additional charge 10 of solid fuel is offset from the front end of the charge 5 of the march stage along the axis towards the march nozzle 8 by a distance of 1 up to 3 outer diameters of the housing 9 of the starting engine. An additional charge of solid fuel is made of LK-6T fuel and is made with the addition of at least 10% metal particles in the form of a fine powder with a mass calorific value of more than 24,000 kJ / kg. The march stage 5 charge is made of DAMST + B boron-containing fuel with a mass calorific value of more than 43,000 kJ / kg (see L.S. Yanovsky, V.N. Aleksandrov et al. Integrated ramjet solid propellant engines. M: ICC "Academbook", 2006, p. 72, Table 2.2 and p. 149, Table 3.8).

Integrated ramjet engine operates as follows.

When the ignition device (not shown) is ignited, the surface of the charge 7 of solid fuel starting stage. The operation of the starting engine is realized at a high working pressure, which leads to high traction and economic performance of the starting stage. During the operation of the starting engine, an increase in the flight speed of the aircraft occurs. In acceleration mode, the input of the intake device 1 is closed with durable resettable plugs. At the end of the operation of the starting solid propellant engine, the aircraft accelerates to the required flight speed to turn on the marching ramjet engine. After that, the air intake device 1 is freed from the plugs at the inlet (not shown), and the sustainer stage is launched. To eliminate the failure of the traction and reduce the speed of the aircraft at the beginning of the march mode, the ignition of the charges of the fuel march stage can be carried out until the start engine is completely finished. In the marching mode, the air coming from the intake device 1 is divided into two parts: the smaller one enters through the supply channel 2 to the solid fuel charge channel 5, and the rest is fed to the combustion chamber 4 through the air bypass channels 3 for the rear end face of the solid fuel march charge 5 steps. The burning of an additional fuel charge 10 at the beginning of the march mode leads to an increase in the hydraulic resistance of the combustion channel and causes a redistribution of air flow between the air supply and bypass channels 2, 3, which improves the ignition of the march fuel in the front of the charge 5. Moreover, the total air flow through the engine decreases, and the specific critical impulse of thrust and engine thrust increase, which ensures the thrust is boosted at the beginning of work, the aircraft is accelerated to the required speed and reliable th air intake device start the engine 1. The degree of boost depends on the additional length 10 of fuel charge for a given thickness thereof, and the wall thickness of the additional fuel charge 10 determines the engine boost. Due to the better ignition of the sustainer fuel in the front of the charge 5, the fuel burns more evenly along the length of the charge 5. To stabilize the combustion process and evenly distribute the air flow in the direct-flow path, a stabilizer 6 is located in front of the front end of the charge 5 of the sustainer fuel. Air entering the charge channel 5 solid fuel march stage, used for primary gasification of fuel. The resulting primary gasification products, enriched with combustible components, then enter the free volume of the combustion chamber 4, where they are mixed with air from the bypass channels 3 and burned out. At the active site, a constant flight speed is maintained due to auto-regulation of fuel consumption. Auto-regulation of fuel consumption is based mainly on the use of a power-law dependence of the rate of gasification of fuel on the gas current density in the channel of the block (L.S. Yanovsky, V.N. Aleksandrov, etc. Integrated ramjet engines using solid fuels. M .: ICC "Academbook", 2006, pp. 161-184).

In the case of manufacturing an additional charge 10 of compositions with a low burning rate of type LK-6T, the additional charge 10 is burned during the entire operation of the marching stage, which provides an additional heat flow to the combustion surface of the charge 5 of the solid combustible marching stage. This leads to the creation of a local zone of increased temperature (more than 2000 K), which provides a significant increase in the completeness of combustion and, accordingly, the specific impulse of the march stage of the propulsion system. Creating a hot zone is especially important when using high-energy boron-containing compounds, for example, DAMST + B, etc. as fuel for the 5th march stage charge. To realize the energy potential of boron and its compounds, it is necessary to ensure that the temperature in the combustion zone is not lower than 1900 K, and the mode will be implemented combustion without an oxide film on the surface of boron particles, which significantly increases the energy release of combustion products and leads to an increase in specific impulse. At lower temperatures, boron and its compounds are coated with an oxide film, and the energy potential of the fuel remains unrealized.

Thus, the proposed IPVRDT allows to improve the flight performance of aircraft.

According to the claimed technical solution, an experimental setup was made with the configuration of the flow part of the integral engine. The conducted fire tests showed that the combustion of the additional charge provided an increase in the completeness of combustion and, accordingly, in the specific impulse of the marching stages with the charge of high-energy boron-containing solid fuel.

Claims (4)

1. An integral ramjet solid-fuel jet engine comprising an air intake device with air supply and bypass channels, a combustion chamber with a channel charge of a solid marching sustainer stage located in front of it, a flame stabilizer, a charge of solid fuel of a launch stage and a marching nozzle, the channels air supply and bypass connect the intake device to the combustion chamber in the front and rear parts of the charge of solid fuel marching stage, characterized in that the charge is solid of the starting stage fuel is made in a cylindrical housing and placed centrally with an annular gap in front of the combustion chamber, and an additional ring charge of solid fuel is installed on the outer surface of the starting stage housing. 2. The engine according to claim 1, characterized in that the outer surface of the additional fuel charge has a taper angle of up to 30 degrees to the axis of the integral engine. 3. The engine according to claim 1, characterized in that the additional fuel charge is offset relative to the front end of the charge of the sustainer stage along the axis towards the sustainer nozzle at a distance of 1 to 3 outer diameters of the housing of the starting engine. 4. The engine according to claim 1, characterized in that the additional charge of solid fuel is made with the addition of at least 10% of metal particles in the form of a fine powder.

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