Air engine design

FIELD: engine engineering, aviation.

SUBSTANCE: in accordance with the present invention fillet fairing is installed inside bypass channel of the external engine circuit in order to avoid outer thickness of auxiliary mechanisms and gear boxes and to actuate them. The external circuit channel is between engine housing and inner circuit of compressor/engine turbine. The fillet fairing dimensions are enough to accommodate auxiliary mechanisms. At the same time the external circuit channel is correspondingly made axisymmentrical to avoid or compensate any blocking effect from fillet fairing within the channel limits when air flows. In addition the fillet fairing may be provided for placing engine oil tank as well as filter/heat exchanger mechanisms foreseen for engine. Under the above circumstances it is essential that elongated cylindrical engine profile is maintained so that reduced cross section is required allowing for the engine to keep a reduced glider of air craft. As a result acoustic shock waves profile is improved.

EFFECT: elimination of outer thickness when auxiliary mechanisms are placed.

 

The present invention relates to developing devices, aircraft engines, and more specifically refers to the layout of the engines used in aircraft and high supersonic flight speeds.

In relatively modern aircraft gas turbine engine some auxiliary units, such as a carton of actuators and electric starter/generator is installed outside of fan housing inside the gondola or glider, within which is embedded the engine. Assistive technology, such as the supply pipeline for lubricating oil and electric cables are laid through the fairings, passing across the channel in the external circuit. These fairings not accept structural power load, but provide aerodynamic profiled shape around the auxiliary technical means.

To minimize aerodynamic drag gondola and glider tightly hug the engine on its periphery, reducing frontal area to its minimum value. However, there is one disadvantage, which is that when profiling the aerodynamic shape of the airframe or gondola through compromise presence of thickening for additional technical means. It is clear that about atanomy any thickening can be given a streamlined shape but this implies that such action will increase the drag coefficient of the aircraft due to the presence of steeper angles of the engine, required to perform a working operation cleaning auxiliary means. It is already known that at supersonic flight speeds of the aircraft such thickening will increase sound blow on the gondola.

In the description of the invention under the patent of great Britain No. 744,695 disclosed compact bypass gas turbine engine includes an internal circuit having consistently placed the compressor, combustor and turbine. The flow core flow is turned and facing forward in order to flow through the combustion chamber, which is placed within a set of discrete nozzles. The engine additionally includes a discrete bypass flow tube, which in the peripheral direction with alternating placed between the tubes in the combustion chamber. Since the tubes of the combustion chamber extends only to the axial area of overflow pipes, assistive technology engine placed between the bypass tube and the front axial section of the tubes of the combustion chamber. Although this configuration of the engine is shortened due to the presence of counter-flow combustion chamber, has rezny lack, namely, that the reversal of the gas flow causes a significant loss of energy flow and interruption of admission of the gas stream in the combustion chamber. In addition, peripheral alternating bypass tubes and tubes of the combustion chamber means that, at any given air flow through the core flow in the engine has not only the annular inlet, but also the core flow countercurrent to the combustion chamber is essentially a substantial plot in the annular channel of the external circuit in the modern conventional gas turbine engine. Thus, the bypass gas flow leads to significant energy flow losses, the energy comes in discrete bypass tube. Thus, the frontal area of the engine can be significantly increased compared with the frontal area of a conventional gas turbine engine having an annular channel in the external circuit in the absence of flow countercurrent to the combustion chamber. Moreover, in the description of the invention under the patent of great Britain No. 744,695 not disclosed nor fairing, passing across the channel in the external circuit, no fastening the auxiliary technical means within such a fairing. The engine, made in accordance with the invention under the patent Great is Britain's No. 744,695, is not suitable for flight at high speeds or supersonic speeds.

In accordance with the present invention provides for the creation of a gas turbine engine containing the axis of rotation, the fan, the inner contour of the engine is surrounded by an outer casing with the formation of the bypass channel, assistive technology engine and fairing, the fairing runs essentially radially between the inner contour of the engine and the outer housing, and the engine is different in that its subsidiary technical means disposed within the fairing.

Assistive technology related to the internal circuit of the engine with the opportunity to receive their drive from it through the use of the drive shaft.

There are some technical AIDS that contain a box of drives, and other AIDS that are mounted on the first with the possibility of obtaining drive from them.

Preferably assistive technology posted by essentially axially with respect to the axis of rotation of the motor, while the other auxiliary tools are installed essentially axially along a box of drives for information to the minimum value of the cross-sectional area of the fairing.

Alternate the VNO assistive technology installed essentially perpendicular to the axis of rotation of the engine.

Under alternative auxiliary technical means are located at an angle relative to the perpendicular to the axis of rotation of the engine and a relatively straight line parallel to the axis of rotation.

Preferably other technical AIDS are made with the same size, which determines the aerodynamic shape of the fairing.

Preferably provides at least two fairings, and thus their location is not required, at least two fairings under normal the annular arrangement of vanes.

Preferably, the fairing is adapted to transfer loads from the engine between the internal motor circuit and outer case, while structural loads contain any one or more loads from a group that includes axial traction, lateral, vertical, or twisting power load. When the fairings are curved and are made so as to straighten return the fulfilled flow of air from the fan.

Preferably the engine is surrounded by a nacelle in order to minimize the value of the aerodynamic resistance.

Preferably, the deflectors are in position aerodynamic trim in the direction across the engine.

Preferably, the about least one case was adapted to normalize the air flow through the bypass channel, and such adaptation is done by making barrel-shaped, at least one case.

Preferably fairings and/or building accessory gearboxes provide thermal protection for mechanisms supporting technical resources.

Under alternative oil tank and/or heat exchangers for diesel fuel placed in the fairings.

Preferably the section of the channel in the external circuit is movable in order to have access to the fairing.

Under alternative a door for access is provided in the housing, the door to access provided for in the fairing.

Variants of the embodiment of the present invention will now be described in disclosure example of its implementation with reference to the accompanying drawings, in which:

figure 1 is a side cross-section schematic of a gas turbine engine mounted inside the nacelle and is known for the current level of technology;

figure 2 schematically depicts a side view of a gas turbine engine when possible installation on the wing for high speed flight;

figure 3 schematically shows a longitudinal cross-section made along the horizontal Central line of the device of aviation (the country : Russia the wow engine in accordance with the present invention;

figa - section of the fairing along the line a-a in figure 3;

4 is a schematically depicts a cross-section bow section of the device of aircraft gas turbine engine in accordance with the present invention;

5 is a schematically shown cross-section of the fairing along the line b-b In figure 4, made in accordance with an additional variant of the present invention;

6 is a schematically depicts a section front section of aircraft gas turbine engine, taken along the horizontal center line of the device of aviation gas-turbine engine, in accordance with the present invention.

Paying attention to figure 1, one can see that shows pos.10 usually known from the existing prior art bypass gas turbine jet engine has a main axis 11 of rotation. The engine 10 includes consistently placed in the axial flow air intake 12, a fan 13 of the engine and the inner contour 8 of the engine, which itself contains the compressor 14 medium pressure compressor 15 high pressure equipment 16 for operation of the combustion chamber, the turbine 17 of the high pressure turbine 18 medium pressure turbine 19 low pressure. The engine 10 includes an exhaust nozzle 20. The gondola 21 essentially surrounds the engine 10 and forms as vozduhozabora is 12, and an exhaust nozzle 20.

Gas turbine engine 10 works in a conventional way and so that the air entering into the air intake 12, would be accelerated by the fan 13 with the aim of obtaining two air flows: a first air flow flows into the inner contour 8 of the engine and through the compressor 14 medium pressure and the second air stream passes through the channel 22 of the external circuit in order to provide axial thrust of the engine. The compressor 14 medium pressure compresses the air stream directed at him, before this air into the compressor 15 high pressure, where it is further compressed stream.

The compressed air is released from the compressor 15 high pressure is directed to the equipment 16 for operation of the combustion chamber where it is mixed with fuel, after which the mixture is ignited. The resulting hot combustion products then expand, and thus before it is released through the exhaust nozzle 20 and before the loss of energy forced to work the turbine 17 of the high pressure turbine 18 medium pressure, and the turbine 19 to the low pressure while providing additional traction engine through the exhaust nozzle 20. The turbine 17 of the high pressure turbine 18 medium pressure turbine 19 low pressure respectively actuate the compressor 15 of the high pressure compressor 14 Wed is the last pressure and the fan 13 via a connecting shaft 23, 24 and 25.

The fan 13 to the periphery is surrounded by a structural element in the form of a housing 41 of the fan, which relies on koltseobrazno installed outlet guide vanes 9 passing between the housing 39, which surrounds the inner contour 8 of the engine.

The engine 10 additionally includes unit 28 boxes of actuator/generator, which is used to start the engine and to generate electricity as soon as the engine starts and continues as normal. The produced electricity is used for the engine and associated aviation support electrically existing technical means well known to the current level of technology. Unit 28 boxes of actuator/generator is kinematically connected with the possibility to drive with the high-pressure shaft 24 by use of a drive means 35. However, in other variants of the embodiment of the invention, the unit 28 of the box actuator/generator can get the drive to any one or more of the shafts 24, 25. In this variant of embodiment of the invention, the unit 28 of the box actuator/generator includes inner box 29 drives, connecting the first drive shaft 30 to the shaft 23 of the high pressure, intermediate box actuators 31 attaching the first lead wall to the second drive shaft 32, and outer carton 33 drives connected to the possibility to drive with the second drive shaft 32. Outer carton 33 drives with the possibility of actuator connected with the generator 34, which has the ability to work on the above-mentioned method. The generator 34 and outer carton 33 actuators are mounted on the casing and disposed within the nacelle 21. The first drive shaft 30, the intermediate box 31 drives and the second drive shaft 32 disposed within the distribution of the fairing 40 channel external circuit.

According to the link on pages 66-71 5-th edition of the book "The Jet Engine", published by the Rolls-Royse pls in 1986, ISBN 0902121235, box 33 drives not only results in the working position the starter and the generator 36, but is also a drive for other auxiliary means, such as a set of pumps. Traditionally, box 33 of the drive and driven auxiliary technical means (36) are located around the periphery of the housing 41 of the fan and in the lower part of the engine 10.

Other auxiliary technical means 36, known from the existing art, is also mounted on the fan case.

In General, gas turbine engine includes multiple rotating blades of compressors 13, 14, 15 and turbine blades 17, 18, 19 located wok is ug common axis 11. In such circumstances, speculative gas turbine engine appears to be cylindrical. Thus, the basic form of a gas turbine engine are longitudinal cylinder and any mechanisms 28, 36 auxiliary means, which will be out in relation to the basic cylindrical shape. As for high speed aircraft, it should be noted that its aerodynamic profile and the boundary of the flight modes are of great importance in relation to the coefficient of resistance, and in relation to sound beats/noise. Under such circumstances, the existence of leading edges and bumps caused by mechanisms boxes actuators and auxiliary means in addition to the main cylindrical profile of the engine, causes problems when seeking to reduce to a minimum the amount of aerodynamic drag.

Figure 2 illustrates a typical device of engine for high-speed flight, located on the wing 2 of the aircraft 3. As can be seen, the wing 2 is connected with the gas turbine engine 10. In the presence of high velocity and at a potential supersonic flight speeds nasal air intake of the nacelle is inconvenient to use because of the severity of problemapomorphine shock wave, and so continually decreasing the efficiency of air intake is manifested by increasing the flow velocity of the inflowing air. Thus, at high flight speeds so-called input configuration when external / internal compression, when the supersonic flow of air, which air intake is significantly reduced to subsonic flow, lead to the fact that they are preferred to align with the needs of the compressor of the engine. This type of device air inlet, as shown in figure 2, creates a series of soft shock waves without excessive reduction of the efficiency of the air intake of the compressor.

In order to reduce the aerodynamic resistance diameter of the fan is maintained at its minimum value, which results in the presence of a relatively large length of the engine. Relatively long and thin profile of the engine 10 is obtained by taking into account compromise by meeting the requirements allocation mechanisms auxiliary technical means within the nacelle 21, which in this example leads to the presence of at least one speaker thickening 5 in the lower part of the engine 10. This thickening 5, although aerodynamically smooth, still increases the drag coefficient, as well as syvaet increased the intensity of the sonic boom.

In the ideal case, the profile engine within the nacelle should be kept to a minimum its characteristics in order to achieve as low a value of the drag coefficient, as well as at high flight speeds to reduce problems associated with noise in the environment if it sonic boom.

The present invention relates to the layout engine, in which the mechanisms auxiliary means are located within the main cylindrical profile of the engine, thus greatly reducing drag and helping to minimize sonic boom.

Turning now to Fig 3 and Fig 4, we can see that supported generally cylindrical profile of the nacelle or housing 21 of the engine 10 at the time, as the mechanisms supporting technical means is placed within this profile. The engine 10 is basically constructed as described with reference to figure 1, but now will be considered the distinctive features described in relation to the present invention.

In accordance with the present invention is provided by the presence of the Radome 26, which is located within the channel 22 of the external circuit, which covers mechanisms 27 auxiliary means. This mechanism is 27 auxiliary means include unit 28 boxes of actuator / generator, as well as other auxiliary technical means 36, such as pumps for pumping oil pumps for fuel, electric generators are designed to supply electricity to facilities for the airframe and hydraulic operating mechanisms. Essentially, box 28 drives now axially paired (axis 11), and each of the driven auxiliary means 36 also essentially axially connected within the fairing 26. Thus, the axis of rotation of the auxiliary means 36 that receives the drive from the box 28 drives, essentially, are perpendicularly oriented relative to the axis 11 of the engine.

Although it is preferred pairing accessory gearboxes and auxiliary means essentially parallel to the axis 11, is also possible for them to mate essentially perpendicular or even at an angle between parallel lines and perpendicular. The advantage of this is that the drive lever 54 communicates with the gear 28 drives in the presence of the giver of the benefit and the desired angle (see figure 3) depending on where the drive lever 54 communicates with the inner contour 8 of the engine and where the box 28 of the actuator is mounted within the cowling 26.

The fairings 26 are located within the General is about the cylindrical profile of the engine 10 and do not create speakers thickening, as described with reference to figure 2. In contrast to the devices known from the prior art, the present invention allows to be closer to the shape of the cylinder profile gondola, which significantly reduces aerodynamic drag and/or diminish the importance of the sonic boom of an aircraft.

Mechanisms 27 auxiliary means are connected in such a way as to ensure their required operation in accordance with known processes.

On figa shows the preferred arrangement of auxiliary technical means 28, 36, located within the fairing, and the profile of the fairing 36. Box 28 drives located inside relative to the support means 36 and is oriented radially. Box 28 drives are connected with the possibility to drive with the inner contour 8 of the engine through the drive shaft 54 and, in fact, paired and arranged axially, due to what is the smallest region in relation to stream flow in the bypass channel. Each auxiliary technical means 36 that receives its drive from the box 28 of the actuator is such that the size of each of the auxiliary means 36 is conveniently AE determines dinamicheski profile of the fairing 26. This arrangement auxiliary means 28, 38 especially brings the advantage of reducing the degree of blocking the flow in the bypass channel 22.

It should be noted that at least one other fairing 26' may be included in the engine, and the fairing includes other auxiliary technical means 27'.

Traditionally many of the guide blades 9, koltseobrazno placed (see figure 2), capable of transmitting wind power loads between the inner circuit 8 and an outer housing 41 of the fan, and then on the mounting structural mount 58 of the aircraft (see figure 4). An additional advantage of the present invention is that the fairings 26, 26' are designed in such a way that they perceive the aerodynamic power load. When implementing the present invention, at least part of the guide vanes 9 may be replaced by collars 26, 26', although there is a possibility that all the ordered set of vanes will be replaced in the case, if it is provided by the presence of not one but more fairings 26, 26'.

In this case (see figure 4 and 6) fairings 26, 26' are rigidly connected between the body 39 of the inner contour of the engine and the housing 41 of the fan or casing 21. The fairings 26, 26' is predstavlyaet a rigid box like structure 60, able to absorb the axial traction, vertical and horizontal power loads, and the twisting force of the load acting on the engine. It should be noted that those skilled in the art may be many different structural forms, but this person will easily understand that such alternative forms must be technical means for transferring power loads acting on the engine, between the inner contour 8 of the engine and the housing 41 of the fan. The fairings 26, 26' so rigidly attached to the outer casing 41 or 21 and to the body 39 of the inner contour of the engine, each case will be finished, essentially annular in shape and it would be characterized by high stiffness. When the fairings 26, 26' are in axial direction at a relatively great length compared to the Radome 40 (see figure 2), known from the current level of technology in its given field, realized additional benefits of increased rigidity of the inner contour of the engine. Such advantages include improved monitoring gaps at the end sections of the blades, thereby increasing the efficiency of the operation.

Turning now to figure 5, it is possible to ensure that well-known is the fact that the guide vanes also provide the Xia flow straightener return the fulfilled air, coming out of the fan 13. In the further development of the present invention to achieve additional benefits fairings 26, 26' are also curved in order to achieve a similar flatness return the fulfilled of the air flow.

Turning now to Fig.6, it can be seen that the present invention allows the placement of mechanisms 27 auxiliary technical means within the Radome 26, but you can understand that the fairing 26, positioned within the bypass channel 22 can cause turbulence, blocking and heterogeneity in flow 24. In such circumstances, the internal formation in the framework of essentially concentric placement of the housing 21 of the nacelle and the housing 39 of the inner contour of the engine for the formation of the bypass channel 22 with the purpose to implement the control stream 24 to achieve efficient operation of the engine 10. Internal formation includes making barrel-shaped in the presence of a concentric arrangement of the housing 21 and the housing 39 of the inner contour of the engine in order to limit the effect of the introduction of the fairings 26 into the channel 22. Giving it the barrel-shaped and includes a radial length of the bypass channel 22 is between the seats, marked radial extensions 44 and 43 and located essentially far from the fairings 26, 26', and consequently places that are directly related to the provisions of the fairings 26, 26'. The magnitude of the radial length 43 is greater than the radial length of 44.

Giving it the barrel-shaped either by forming the housing 21 of the external circuit or housing 39 of the inner contour of the engine, or the giving of bacholrette as the housing 21 and the housing 39. When the fairings 26, 26' change width on its periphery in the direction of the flow due to the change of the size of the auxiliary means, placed in these fairings, the degree of bacholrette form is also changed to maintain a constant or desired in other cases, the cross-sectional profile of the air flow. Note that the degree of giving bacholrette forms must be relatively small in terms of value, and that the outer profile of the nacelle must be supported in a cylindrical shape, as described above, indicating the benefits received from this profiling.

You should also understand that although there is a possible increased harmful effects of auxiliary means to the air stream 24 and the traditional orientation of the auxiliary materials is lnyh technical means (in the horizontal direction on the Central line), it is customary to have the ability to perform three fairings, which are located at angles of 120 degrees, or even four fairings set at angles of 90 degrees. Under alternative fairings in accordance with the implementation of the present invention can be in an unbalanced position within the block cross-section when such asymmetric changes caused by the change of the cross section of the bypass channel, or for other reasons.

In addition to the fairings 26, which are mechanisms 27 auxiliary means, it should also be understood (see figure 4)that can be included fairings 26, which are used simply as tanks 34 for content of lubricating oil, or they may contain oil filters 35, or they can be designed to ensure the appropriate location of the heat exchangers 45 for cooling oil or fuel. If any features fast maintenance it is advisable to place them near the designated panels with cutouts or panels 50 to provide access, located in the buildings of the engine.

Note that the inner contour 8 of the engine, including the combustor 16, and a turbine 17, 18, 19 and other devices during operation of the engine becomes relatively nagretye such circumstances, the fairings 26 contain the appropriate shielding means 31, 52 mechanisms for 27 additional technical means of the engine, protecting them from exposure to temperatures internal contra 8 engine. In one variation of the embodiment of the present invention this is achieved by use of the casing 31 of the box actuator and the housing seal 39 of the inner contour of the engine and fairings 26 in order to escape assistive technology in a separate area. However, it should be considered that the flow of air through the channels 23 will provide cooling cowl 26, and this, in turn, should limit the influence of nagrevaemoi in relation to mechanisms of auxiliary technical means held within the fairing 26.

In General, mechanisms 27 held within fairings 26 will be supplied with energy taken from the energy that is produced in the inner contour 8 of the engine and spend traction movement through adjacent box 28 drives. Thus, the corresponding radial actuators 54 (see figure 4)working with the inner contour 8 of the engine, give energy to drive these boxes 28 of the actuator and, thus, mechanisms 27 auxiliary means, located within the fairing 26. Alternative each of the auxiliary means 27 may soon get an individual the actual drive from the electric motor 56, than the radial actuator, powered by the engine.

Operation of gas turbine engine 10 can be carried out in accordance with traditional technology of its operation except that the fairings 26 allow the placement of mechanisms 27 auxiliary technical means within the traditional profile of the fairing of the engine 10. In short, mechanisms 27 auxiliary means are located within the fairing 26, which pass through the width of the bypass channel 22. Air flow 24 is supported by the corresponding unbalanced formation, and to provide a barrel-shaped channel 23 of the external circuit to mitigate the effects of blocking caused by the fairing 26. Under such circumstances, even so, make bacholrette the housing 21 of the engine 10 has a reduced diameter compared to those who have known the engine when implementing other appropriate considerations (for example, when there are deflected profile of the blades of the fan or in the presence of a piping system or system of channels of the engine, placed between the fairing 37 gondola and the housing 21 of the engine). This diameter dictates the minimal size of the gondola.

It is clear that it is necessary to carry out technical maintenance of the carton 28 drives and mechanisms 27 auxiliary means held within the fairing 26. In such circumstances, access to these fairings 26 and mechanisms 27 through intended for this sash 50 access. These shutters 50 access are located within housing 21 and form part of its design, and the housing 21 defines a channel 23 in the external circuit of the engine. Sash 50 is made in the form of articulated sections of the channels, swivel mounted relative to the housing 21 of the gondola. Under alternative sash 50 access can be removed. Sash 50 access provide improved rigidity of the channel 23 during the flight, while the fixed section 41A of the housing 41 to provide structural strength to maintain the parts of the engine (for example, Assembly of the thrust reversing/adjustable nozzle). Sash 50 provide access during the performance of maintenance operations to the auxiliary technical means, located in the fairing 26, as well as access to parts of the inner contour 8 of the engine.

Paying attention to Fig.6, we can see that alternative access includes the movable section 21A of the nacelle 21, the slide 62 of the housing 41, the slide 64 fairing 26 for access to the auxiliary technical means 27 and p is movable partition panel 66 of the housing 39 of the inner contour 8 of the engine. Although in the description of the invention is indicated that all these panels 21A, 41, 62, 64 access can be movable, they can be mounted swivel or be able dismantling and fitting through mechanisms known from the current level of technology in this field.

To implement a more desirable profile of the engine 10, and hence to perform a more desirable cross-section of the airframe or gondola, in which the engine 10 will be located, it should be considered that can be improved characteristic sound of impact as compared with that found in traditional devices of a gas turbine aircraft engine, designed for flight at higher speeds. In addition, the elimination of harmful aerodynamic effects of external thickening, causing increased resistance fairing or body, should improve performance aircraft. In addition, if there is any thickening to obtain uniformity of air flow, it should be distributed more in the lateral than in the vertical direction, i.e. across the airframe, fuselage or wing. More correct profile of the engine 10 that is compatible with the basis of cylindrical form, allows to reduce the required cross-sectional area of the nacelle 21, formed around dvigatelya, that, in turn, will allow us to determine the profile of the fuselage of the glider within the accepted norms of aircraft design, but with a progressive diminution of the intensity of the sound blows upon receipt of specific benefits to supersonic flight.

While in the foregoing description of the invention attempt to draw attention to those features of the invention, which, as I am sure the author of the invention, are having particular importance, it should also be understood that the applicant claims protection in respect of any patentable feature or combination of features, which previously made the link, and here the above-mentioned and/or shown in the accompanying drawings, regardless of whether it was emphasized specifically with a particular emphasis.

1. Gas turbine engine containing the axis of rotation, the fan, the inner loop, surrounded by an outer casing with the formation of the bypass channel, assistive technology engine and fairing passing essentially radially between the inner contour of the engine and an outer housing, assistive technology engine placed inside of the fairing, where AIDS contain a box of drives, and other auxiliary technical means are mounted on at azannyh support tools with the possibility of obtaining from them the actuator, characterized in that the other auxiliary technical means mounted in axial sequence to the axis of rotation of the engine along a box of drives for the formation of the minimum cross-sectional area of the fairing.

2. Gas turbine engine according to claim 1, characterized in that the auxiliary technical means connected with the actuator with the internal contour of the engine through the drive shaft.

3. Gas turbine engine according to claim 1, characterized in that the other auxiliary technical means have dimensions that define the airfoil shape of the fairing.

4. Gas turbine engine according to claims 1 to 3, characterized in that it includes at least two fairing.

5. Gas turbine engine according to claim 4, characterized in that the fairing is adapted to transfer loads from the engine between the internal motor circuit and outer case.

6. Gas turbine engine according to claim 5, characterized in that the structural loads include any one or more loads from a group that includes axial traction, lateral, vertical, or twisting power load.

7. Gas turbine engine according to claim 6, characterized in that the deflector is curved and designed to straighten return the fulfilled the flow of air is from the fan.

8. Gas turbine engine according to claim 7, characterized in that the engine is surrounded by a nacelle in order to minimize the value of the aerodynamic resistance.

9. Gas turbine engine according to claim 1, characterized in that the fairings are aerodynamically balanced in the direction across the engine.

10. Gas turbine engine according to claim 1, characterized in that at least one housing adapted to normalize the air flow through the bypass channel.

11. Gas turbine engine of claim 10, characterized in that such adaptation is done by making barrel-shaped, at least one case.

12. Gas turbine engine according to claim 1, characterized in that the fairing and/or the body of the box drives provide thermal protection for mechanisms supporting technical resources.

13. Gas turbine engine according to claim 1, characterized in that the fairing is placed oil tank and/or heat exchangers diesel fuel.

14. Gas turbine engine according to claim 1, characterized in that the area of the overflow channel is movable to provide access to the fairing.

15. Gas turbine engine according to claim 1, characterized in that the casing has a door access.

16. Gas turbine engine according to claim 1, characterized in that the fairing is provided by the shutter access.

17. The flyer is the first apparatus, includes gas turbine engine as described in any of the preceding paragraphs.



 

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1 dwg

FIELD: mechanical engineering; gas-turbine engines.

SUBSTANCE: invention relates to units of drives of gas-turbine engines of aircraft and ground application. Proposed device contains fastening members and movable member telescopically connected with casing. Movable member is on accessory gear box and secured by pressure flange for displacement along end face of box. Clearance is formed between mating surfaces of flange and movable member.

EFFECT: improved reliability of connection of casing and accessory gear box by providing their rigid connection and sealing of inner spaces.

3 dwg

FIELD: engines and pumps.

SUBSTANCE: invention refers to aviation and particularly to devices for restraint and arrangement of auxiliary equipment in turbojet engines. The device consists of two coaxial rings (12, 14) assembled one into another and connected to each other with hollow radial poles (16, 18, 20 and 22). The pipelines and electrical wires run inside poles. At least one of the side poles (16, 18) bears a removable panel (24, 26) on its side, which after dismounting facilitates an access to the equipment of the turbojet engine arranged radially inside the interior ring (12) in one line with the radial pole (16, 18). Such design of the device provides an access to the equipment, installed in a turbojet engine.

EFFECT: facilitating access to equipment, assembled in a turbojet engine due to arrangement of a restraining device and of auxiliary equipment.

10 cl, 5 dwg

FIELD: electricity.

SUBSTANCE: cable bundle (3) positioning and retaining station on the turbojet engine (1) housing (2) includes bundle supports (3) standardized for directions perpendicular to the turbojet engine (1) axis (4) and bundle supports standardised for directions parallel to the turbojet engine (1) axis (4).

EFFECT: reduction of manufacturing cycle cost and time.

14 cl, 10 dwg

FIELD: engines and pumps.

SUBSTANCE: invention is intended for feeding electric power to equipment from gas turbine engine. The proposed system comprises an electronic control device to control, at least, one parameter containing the data on originating variation in consumed power, a control valve controlled by aforesaid system and feeding air take off the engine operated in transient conditions and a pneumatic device receiving aforesaid taken-off air to actuate the aircraft onboard equipment. The latter can represent an air turbine or generator with built-in pneumatic circuitry.

EFFECT: use of engine pneumatic power to drive aircraft onboard equipment.

33 cl, 10 dwg

Air engine design // 2355902

FIELD: engine engineering, aviation.

SUBSTANCE: in accordance with the present invention fillet fairing is installed inside bypass channel of the external engine circuit in order to avoid outer thickness of auxiliary mechanisms and gear boxes and to actuate them. The external circuit channel is between engine housing and inner circuit of compressor/engine turbine. The fillet fairing dimensions are enough to accommodate auxiliary mechanisms. At the same time the external circuit channel is correspondingly made axisymmentrical to avoid or compensate any blocking effect from fillet fairing within the channel limits when air flows. In addition the fillet fairing may be provided for placing engine oil tank as well as filter/heat exchanger mechanisms foreseen for engine. Under the above circumstances it is essential that elongated cylindrical engine profile is maintained so that reduced cross section is required allowing for the engine to keep a reduced glider of air craft. As a result acoustic shock waves profile is improved.

EFFECT: elimination of outer thickness when auxiliary mechanisms are placed.

FIELD: engines and pumps.

SUBSTANCE: proposed unit consists of gas turbine and reduction gear accommodated inside container and coupled via transfer shaft, reduction gear output shaft carrying the pump. Input device is arranged between said reduction gear and engine so that device front face wall part seats on reduction gear, while device read face wall part is located on gas turbine engine. Note here that both aforesaid parts are linked up axially and radially by sealed telescopic couplings with the remaining part of input device fixed container. Inlet inspection window is made in input device front face wall. Input device lower wall is made flat and horizontal. In operation, sealed telescopic couplings allow the engine and reduction gear to move relative to input device with no loss in tightness on the latter.

EFFECT: higher reliability, reduced weight and overall dimensions, easier mounting and control.

3 dwg

FIELD: machine building.

SUBSTANCE: unit consists of gear box of gas turbine and of at least one starter/generator mechanically coupled with gear box. The gear box consists of gears with several pinions. The starter/generator contains a generating block with a rotor, forming an inductance coil and stator forming an anchor; further, the stator/generator contains an actuating block with the stator forming the inductance coil and rotor forming an anchor connected to the inductance coil of the generating block. The rotor of the generating block and the rotor of the actuating block are arranged on a common shaft with a pinion engaging the gear of the gear box on both sides of this pinion. The invention facilitates integration of the starter/generator into the gear box of the gas turbine.

EFFECT: reduced volume and dimensions, ensuring easy disassembly.

23 cl, 8 dwg

FIELD: engines and pumps.

SUBSTANCE: auxiliary mechanism drive of two-shaft gas turbine engine comprising high- and low-pressure shafts incorporates first mechanical transmission between high-pressure shaft and drive box, and hydraulic transmission between low-pressure shaft and drive box. Auxiliary mechanisms are arranged in drive box, while hydraulic transmission is mounted to allow auxiliary mechanism drive rpm being equal to high-pressure shaft rpm.

EFFECT: possibility to take off power from high- and low-pressure shafts without varying auxiliary mechanism rpm.

14 cl, 3 dwg

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