RU2716651C2 - Double-flow turbojet engine nozzles system - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: invention relates to aircraft engine building, particularly, to double-flow turbojet engines nozzles. System of nozzles for working gases of double-flow turbojet engine without mixing of circuits flows comprises two cold channels changing their cross-section from semi-ring behind engine fan to rectangle at outlet and one hot channel of round cross section located between cold channels. Channels terminate by nozzles, which are devices for control of value and direction of working gas jet. Invention makes it possible to fit nozzle into tail part of aircraft plane by flying wing or tailless and control thrust vector in cold channel, leaving hot channel in form of extension pipe of minimum weight. System of nozzles for working gases produced by gas generator plant of double-flow turbojet engine in aircraft includes channels and nozzles. Channels of cold circuit discharge air flow from engine fan and are formed by cross-sections having shape of semi-ring at output from cold external circuit of engine, with smooth transition of section shape to rectangle located by wide side along rear edge of aircraft wing. Upper and lower edges of such rectangle contain flaps varying the cross-section area on the nozzle edge and/or the direction of air outlet from it. Channel of hot inner circuit removes hot gases from engine gas generator and is formed in cross sections by circles. Longitudinal axis of channel of hot circuit is made in the form of smooth transition from axis of engine to axis parallel to plane of symmetry of aircraft or forming with such axis small angle to the outside in order to reduce turning moment at failure of engine.

EFFECT: invention allows reducing aerodynamic resistance and weight of aircraft.

6 cl, 16 dwg

Description

The invention relates to the field of aircraft engine manufacturing, and in particular to the design of nozzles of dual-circuit turbojet engines (turbojet engines).

A similar transition of the cross-sectional shape from the half-ring behind the fan of the turbojet engine to the rectangular cross-section of the nozzles on the sides of the hot inner loop nozzle is described in US Pat. No. 3,137,131 A, where three nozzles form a vertical gap in order to reduce noise on the ground from a running engine. The transition of the shape of the cross section of the jet from a round or annular one at the exit of the turbojet engine to a rectangular one inscribed in the trailing edge of the wing or tail, ending with a flat nozzle that is adjustable in area and direction of exit of the jet of gases, in order to combine the functions of the gas-dynamic and aerodynamic controls and reduce the resistance of the tail section protected by many patents, however, the combination of a round nozzle of a hot inner contour with a flat one located along the trailing edge of the wing or plumage and equipped with means the author does not know the speed and direction of flow of the nozzle flaps of the external cold circuit of the turbojet engine and is the subject of this invention. Known flat nozzle turbojet engine containing a manifold that separates the gas flow into two half nozzles, and means for adjusting the magnitude and direction of flow (see RF patent No. 2443891, CL F02K 1/40, publ. In 2012), which is the closest analogue of this application according to the stated goals.

A disadvantage of the known device is that the collector, when used with a turbofan engine, is made with a mixture of a cold air stream from a fan (external circuit of a turbojet engine) and hot gases of a gas generator (internal circuit), as a result, a channel of complex shape is blown by a stream of hot gases and must be made of heat-resistant materials, such as steel, which increases its weight. For a turbojet center with a small bypass ratio (the ratio of gas flow through the external and internal circuit), this solution is expedient and leads to a certain decrease in the temperature of the collector walls. On heavy transport aircraft, a high and ultra-high bypass ratio is used to reduce fuel consumption. When using the aerodynamic scheme, a flying wing or tailless for a heavy transport aircraft, the placement of the turbofan engine above the fuselage is complicated by the large diving moment of the engine thrust force, which complicates the separation of the nose wheel during take-off. The transfer of engines to the rear edge of the center section and the rear of the fuselage unacceptably shifts the center of gravity of the aircraft back. Therefore, layout schemes are possible with the placement of the turbofan engine in the thickness of the center section, which reduces the area of the washed surface of the aircraft, but leads to an increase in the size of the channel that takes air and hot gases from the turbofan engine. Increasing the length of the channel allows you to smoothly change its cross-sections to rectangular and make the nozzle in the form of rotary horizontal flaps, providing adjustment of the cross-section, deviation of the thrust vector and reverse thrust.

For aircraft driven by a turbojet engine, both manned and unmanned, the criterion for perfection is the ratio of the mass of the cargo transported at a given distance to the maximum take-off mass. The fuel mass is reduced by using a turbofan engine with a high bypass ratio and an aerodynamic design of the flying wing, as well as a decrease in the area of the washed surface of the aircraft due to the placement of the turbofan engine in the wing center section. The weight of the structure can be reduced by using thrust vector deviation for balancing and controlling the aircraft instead of part of the steering surfaces, using lighter materials for the walls of the cold circuit channel of the turbofan engine, and also by including the walls of the cold circuit channel in the power circuit of the aircraft tail section.

In addition, the author of the present invention used a double nozzle in the cold circuit, called a bifurcated, in which the air stream leaving the fan circuit of the turbofan engine is divided into two flows and directed to two nozzles. Both flows are thrown out parallel to the thrust axis. The advantage of this design is that it allows you to control the aircraft, in particular by pitch, roll and yaw, by controlling the flaps either by deflecting the thrust vector or by changing the flow rate. Being at a distance from each other, the nozzles are also offset relative to the axis of the gas generator.

The technical result of the present invention is the creation of a nozzle system that improves the perfection of the aircraft as a whole by increasing the aerodynamic quality, reducing the mass of the structure, replacing part of the aerodynamic controls with gas-dynamic.

In this regard, the object of the present invention is a nozzle system for a turbojet engine in an aircraft (in the example of a flying wing or tailless design in the form of a heavy transport aircraft), driven by hot gases emitted by a gas generator, containing channels and nozzles, characterized in that the channels of the cold circuit divert air flow from the engine fan and are formed by sections having the shape of a half ring at the outlet of the cold external circuit of the engine, with a smooth transition of the cross-sectional shape si to a rectangle located with the wide side along the trailing edge of the wing of the aircraft. The upper and lower edges of such a rectangle contain flaps that change the cross-sectional area at the nozzle exit and / or the direction of air exit from it. It is advisable that the cross section of the channels of the cold circuit in the tail of their length has an elongated shape in the transverse direction with a ratio of width to height of not more than 2.5. A further increase in this ratio leads to a sharp increase in thrust losses in the nozzle.

The channel of the hot inner circuit removes hot gases from the gas generator of the engine and is formed in cross sections by circles. The longitudinal axis of the channel of the hot circuit is made in the form of a smooth transition from the axis of the engine to the axis parallel to the plane of symmetry of the aircraft or forming a small angle outward with such an axis in order to reduce the turning moment when the engine fails.

The invention also relates to the possibility of controlling the flight of an aircraft provided by this last type of gas discharge into two nozzles. The use of the aerodynamic scheme of a flying wing or transition from it to a normal one, with a large center section between the payload compartment and the wing consoles, without tail feathering, is associated with the need to deviate the thrust vector. An object of the present invention is also to provide an aircraft control device, in particular in pitch and yaw, which is effective at low flight speeds.

The device must continuously provide vector changes of small amplitude, without leading to a decrease in the efficiency of the gas generator.

It should be able to provide significant thrust vector deviation for aircraft control needs.

It is advisable that the nozzle system was configured to separate the air flow from the fan into the first and second flows for their release through the first and second nozzles, and contain at least one of the following two control means: means for distributing the flow in each of the two nozzles and means for orienting the thrust vector produced by each of the two nozzles.

Under the nozzle in this application should be understood gas outlet jet nozzle, which enters part of the stream at the outlet of the engine. This term is not associated with any particular form. The presence of two streams in this case is used for separate control of two vectors of thrust components in absolute value and in direction.

Preferably, said nozzles are configured to orient the yaw thrust vector. Due to this, the absence of a keel is compensated or a decrease in its size is possible.

It is also preferred that said nozzles be used for pitch or roll control, or that the nozzle system may comprise two pairs of nozzles, for example, one pair for yaw orientation and the other for pitch orientation. Other design options or combinations thereof are also possible.

The following is a more detailed description of the invention with reference to the accompanying drawings, including:

FIG. 1 is a top view of an example aircraft in which the present invention is applied.

FIG. 2 - typical turbofan engine with a large degree of bypass and separate flow outflow.

FIG. 3 - turbofan engine according to FIG. 2 connected to the nozzle system of the present invention, isometric

FIG. 4 is a horizontal section through the engine and nozzle system.

FIG. 5 is a rear view of a twin engine aircraft in accordance with an embodiment of the present invention.

FIG. 6 is a vertical section along the axis of the channel of the air intake, engine and nozzle of an example aircraft in which the present invention is applied.

FIG. 7 is a section through a plane, of the normal axis of the channel, before entering the engine.

FIG. 8 is a section by a plane, the normal axis of the channel, at the exit of the fan, where the cold circuit is divided into two half rings.

FIG. 9 - 11 - sequence of sections of the transition element of the channel, normal to its axis, towards the exit.

FIG. 12 is a schematic view of an arrangement of nozzles of controls in accordance with the present invention, the full opening position of the nozzle.

FIG. 13 is a schematic view of an arrangement of nozzles of controls in accordance with the present invention, the position of the minimum opening of the nozzle.

FIG. 14 is a schematic view of an arrangement in the nozzle of controls in accordance with the present invention, a position of deviation of the thrust vector.

FIG. 15 is a schematic view of an arrangement of nozzles of controls in accordance with the present invention, the position of a thrust reverse.

FIG. 16 - An example of creating a yaw control moment.

In FIG. 1 shows a non-limiting example of aircraft 1. It is designed according to the “flying wing” scheme, contains a payload compartment 2, two wing consoles 3 and 4, and is driven by two turbofan engines 5 located in the left 6 and right 7 large relative-wing section thickness. Air from the air intakes at the leading edge of the center section is led to the engines through channels 8. The upper skin above the engines and nozzles is not conventionally shown.

In FIG. 2 shows a section in the plane of symmetry of a dual-circuit turbojet engine 5. Air flow enters the engine inlet through the air intake and channel 8 and is compressed in a low-pressure compressor (fan). Further, a part of the flow is compressed in the high-pressure compressor 16, is supplied to the combustion chamber 17, where energy and temperature are increased due to the combustion of fuel. Next, the hot gas-air mixture spins high and low pressure turbines 18, where part of the hot gas energy is mechanically drawn to rotate the compressors, and exits the engine into the inlet tubular element of the channel 15 hot circuit. The turbine rotates the compressor 16 and the cold circuit fan 19, which increases the air speed the flow going to the intake manifold of the cold circuit of channel 10. The engine thrust is created by increasing the flow rate. In the practice of aircraft engine manufacturing, two types of jet nozzles are known: with mixing of flows of hot and cold circuits, as in the opposed invention of the Russian Federation No. 2443891, and with separate outflows characteristic of turbojet engines with a high degree of bypass (ratio of the mass flow rate of cold and hot flow).

In FIG. Figure 3 shows the essence of this application - the conversion of a cold stream of a circular shape at the exit of a cold engine circuit into two flat flows at the trailing edge of an airplane wing of a flying wing circuit in order to reduce aerodynamic drag and create control moments for the deflection of these flows. The cold circuit channel 9 utilizes this design feature when being bifurcated. It distributes the stream exiting channel 10 at the input of two streams in two channels 11 and 12, which end with two rectangular nozzles 13 and 14. Channels 10, 11 and 12 have a shape corresponding to the task of ensuring the separation of the flow into two streams, as well as the transition from the shape of a half ring to a shape with a rectangular cross section. The nozzles have movable upper and lower flaps 22, regulating the size and direction of the air flow.

In FIG. 4 shows a horizontal section along the engine and nozzle. The axis of the nozzle is not parallel to the direction of flight, the nozzle is turned outward to reduce disturbing moments when one engine fails in an aircraft with two engines. The axis of the engine 5 is turned outward from the axis of symmetry of the aircraft to reduce the total angle of rotation of the flow from the air intake to the nozzle, which reduces losses in the gas-air path.

In FIG. 5 shows a rear view of the aircraft — as an embodiment of a heavy transport aircraft, the flying wing 1 scheme, with flat nozzles 13 and 14 of the cold circuit located along the trailing edge of the wing and round nozzles 15 of the hot circuit

In FIG. 6 shows a section in a vertical plane along the axis of the air intake, channel 8, engine 5 and channel of the hot loop 15. Small engine parts for a given image scale are not shown. The axis of the engine is located so that the engine does not protrude above the upper surface of the wing, and the channel of the hot circuit 15 exits to the trailing edge of the wing without large turns of the flow. The engine 5 from the outlet of the cold circuit 10 to the exit of the hot circuit 15 is closed by a casing 21.

In FIG. 7 shows a section in the plane of entry into the engine. A fairing 20 on the lower surface of the aircraft is built around a box of engine assemblies.

In FIG. 8-11 show the geometry of channel 9 in accordance with the present invention.

This channel has a section with an inlet tubular element 10 on the side of the engine fan connected to nozzles 13 and 14. The inlet tubular element of the channel of the hot circuit 15 is directly connected to the output of the engine turbine. As shown in FIG. 8-11, its sections perpendicular to the flow are circles.

The cold circuit channel distributes the stream leaving the channel of the engine 10 at the inlet to two streams in two channels 11 and 12, which end with two rectangular nozzles 13 and 14. According to the presented embodiment, the nozzle comprises two movable horizontal flaps 22, each of which forms a point. This type of air discharge is applied to the aircraft shown in FIG. 1, and is intended to provide adjustment of the nozzle exit section corresponding to the engine operation mode, deviation of the thrust vector up or down to create a pitch control moment, reverse thrust, and if the nozzle or left and right engine thrust vector deviates in different directions, the roll moment. The yaw moment is created by different cross-sectional areas that regulate the air flow and the thrust force of the right and left nozzles.

The shape of the channel gradually changes, starting from the input plane, towards the exit, expanding in the transverse direction and decreasing in height.

The values of the channel area in cross sections are determined depending on the requirements associated with the dynamics of the fluid, to minimize traction loss.

The following is a description of an embodiment of the shutter control system of the nozzle 22 shown in FIG. 12-15. According to this embodiment, the flaps have an axis of rotation 23 and control rods 24 located on the side of the flow part. The control rods are attached to the rocker 25, which rotates on the rod of the cylinder for controlling the opening of the nozzle 26. A symmetric change in the opening of the right and left nozzles depends on the operating mode of the engine. An asymmetric change in the opening of the right and left nozzles is a means of distributing air flow to create a control moment around a vertical axis. A thrust 27 is also attached to the rocker, the front hinge of which is connected to the rod of the control cylinder by a thrust vector 28 fixed to the cylinder rod 26. The reverse flaps 29 are also fixed on the rotation axes 23, the front edges of which are held by locks attached to the fixed part of the nozzle 30. The leading edges of the leaves 22 are equipped with seals to prevent air leakage between them and the fixed part of the nozzle. The extreme rearward position of the rocker 25 corresponds to the maximum opening of the nozzle and is shown in FIG. 12. In FIG. 13 shows the nozzle in the minimum aperture position when the leading edges of the flaps 22 are close to the flaps of the reverse 29. As the cylinder 26 moves further forward, the locks of the flaps of the reverse 29 open and together with the flaps 22 rotate to the position shown in FIG. 15, completely closing the exit from the nozzle back and redirecting the entire air flow to reverse thrust. To change the direction of action of the thrust, the rod of the thrust vector control cylinder 28 is displaced from the neutral position, causing the rocker 25 to rotate and asymmetric deviation of the upper and lower sash 22, which is a means of orientation of the thrust vector. In FIG. 14 shows the deviation of the thrust vector, creating a pitch moment for pitching. To create a control yaw moment without changing the air flow through the engines, create different openings of the right and left nozzles by cylinders 26. For this, a stream separation into two separate streams in a bifurcated exhaust manifold is also used in accordance with the present invention. According to a distinguishing feature of the invention, control means are provided for controlling both aircraft in order to control the aircraft 1 without tail unit. In FIG. 16 shows how a change in the ratio of air flow through the nozzle creates a yaw moment at the center of mass of the aircraft 29.

Preferably, the channels are optimized so that, if the thrust vector is not changed, they create a minimum transverse thrust component of each nozzle. Indeed, this component leads to a decrease in axial thrust, which must be minimized. The total lateral component of the thrust remains zero due to the symmetry of the system. Thus, this way of carrying out work, only slightly affecting the characteristics of the engine, provides the creation of a vector thrust that makes it possible to compensate for the absence of tail empennage, in particular, at low speed modes in transition phases of flight, as well as when the aircraft moves on the ground. As a result, a turning moment is created relative to the center of gravity of the aircraft. This mode of operation provides significant vector traction, allowing you to control the aircraft, however, with some damage to the characteristics of the gas generator. However, this decline is controllable.

The technical result of this invention, in comparison with the traditional arrangement of engines with a large degree of bypass under or above the wing on the pylons on an airplane designed to transport a given mass of cargo at a given range, is:

- The growth of the aerodynamic quality of the aircraft due to the exclusion of the washed surface and the resistance of the nacelles and their pylons

- The increase in the aerodynamic quality of the aircraft by reducing the bottom drag with the ejection effect of the jet engine

- Decrease in the mass of deflected aerodynamic surfaces and their drives due to their partial replacement by nozzle flaps deflecting the thrust vector

- Reduction in aircraft mass due to the combination of thrust reversal flaps with thrust vector deflection flaps in one unit

- Reducing the mass of the aircraft due to the lack of variable in engine thrust pitch moments and aerodynamic control to compensate for such moments.

- Reducing the mass of the aircraft through the use of channels of the cold circuit as the power beams of the tail of the aircraft

Compared with the closest analogue according to the patent of the Russian Federation No. 2443891, class. F02K 1/40, publ. in 2012, which involves mixing the flow of hot and cold circuits and their subsequent separation into two nozzles, the technical result is to reduce the mass of the structure when it is divided into an extension heat pipe - a hot circuit nozzle subject to high thermal loads (about 440 degrees Celsius) and therefore not included in the power circuit of the tail of the aircraft, and two channels of the cold circuit, in which the air after the fan heats up no more than 50 degrees above the ambient temperature. Such channels are the power beams of the tail of the aircraft, perceiving thrust, the moment when the thrust vector is deflected, and local aerodynamic loads. A hot loop nozzle with the possibility of thermal expansion is also fixed on these channels. Cold circuit channels can be made of non-heat-resistant composite materials of minimum weight.

The decrease in the mass of the aircraft’s structure and the increase in its aerodynamic quality, as the technical results of the invention, should exceed the influence of thrust loss in longer air channels of complex shape in take-off weight, then the application of the invention is economically feasible, since the cost of the aircraft is proportional to the take-off mass, at fixed range and mass of cargo.

This application describes an embodiment of the invention. However, other options may be envisaged without departing from the scope of the invention.

Analysis of the totality of all the essential features of the proposed invention proves that the exclusion of at least one of them makes it impossible to fully ensure the achieved technical result.

The analysis of the prior art shows that it is unknown such a device that has inherent features identical to all the essential features of this technical solution, which indicates its unknownness and, therefore, novelty.

The above also proves the conformity of the claimed device to the criterion of inventive step.

When carrying out the invention, the presence of the proposed object is really realized, which indicates industrial applicability.

Claims (11)

1. A system of nozzles for working gases in an aircraft driven by hot working gases produced by at least one double-circuit gas generating unit containing cold and hot channels (11, 12, 15) and nozzles (13, 14), characterized in that the said channels form a gas stream flattened with a wide side in the direction of the trailing edge of the wing, divided into a stream of hot gases from the internal circuit of the engine and two flows of cold air in two cold channels extending from the cold fan circuit engine, with: - the hot channel is placed between the cold and the cross section of the hot channel, normal to the direction of flow, have the shape of circles - the axes of all channels are parallel, - each cold channel contains the first element (10) of the channel in the form of a half ring, the second transition element (11, 12) of the channel, which ensures the transition of the shape of its sections from the shape of the half ring to the rectangular shape of one of the nozzles (13, 14), each of which is made rectangular and is located on the wide side along the trailing edge of the wing of the aircraft on one side of the hot channel extending to the trailing edge of the wing, - the nozzle system contains at least one of the following two control means: means for controlling the amount of air flow in each of the two rectangular nozzles and means for controlling the direction of the thrust vector produced by each of the two rectangular nozzles. 2. The nozzle system according to claim 1, characterized in that said two control means are mechanical means (22). 3. The nozzle system according to claim 1, characterized in that the said rectangular nozzles are configured to create a yaw control torque by redistributing the air flow between the nozzles. 4. The nozzle system according to claim 1, characterized in that said rectangular nozzle is configured to control pitch flight by symmetrically deflecting the flaps up or down, which simultaneously deflects the air flow in the cold engine circuit. 5. The nozzle system according to claim 1, characterized in that the said rectangular nozzles are adapted to control roll flight by opposing the flaps up and down, which simultaneously deflects the air flow in the cold engine circuit. 6. The nozzle system according to claim 1, characterized in that the gas stream is generated by two gas generating units, the nozzle system of each engine comprising means for orienting the thrust vector produced by each of the two rectangular nozzles of the cold circuit of each engine.

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