DE202021104007U1 - Jet engine and jet engine assembly - Google Patents

Abstract

Jet engine (100), in particular designed as a turbofan engine, comprising an air inlet area (3), a housing (7), an engine core area (9) and an inlet cone (17) in the air inlet area (3), the inlet cone (17) having a non-rotatable Component, in particular connected to a guide grille (19), characterized in that the inlet cone (17) is designed as an oblique inlet cone (17), the cone tip (25) being arranged eccentrically to the axis of rotation (23) of the jet engine (100).

Description

The present invention relates to a jet engine comprising an air inlet area, a housing, an engine core area and an inlet cone in the air inlet area according to the preamble of claim 1. Furthermore, the present invention relates to a jet engine arrangement comprising a jet engine according to the invention, an air inlet and a flow channel according to claim 5.

Jet engines in aircraft are exposed to numerous different flow conditions, including during take-off, landing and in flight. In the flight phase, different flight maneuvers can be carried out, for example when changing course. Further requirements arise, particularly at high flight speeds in the supersonic range. A direct flow against blades and stators in the jet engine must be avoided in the case of a supersonic flow, as this would lead to a stall on the blades and stators, suffocate the combustion chamber and the entire jet engine could fail. For this reason, jet engines of aircraft that at least partially fly at supersonic speed are preceded by so-called diffusers as flow channels that slow the supersonic flow to subsonic speeds before the air enters the jet engine. Numerous changes in flow directions and flow conditions occur in these diffusers. Depending on the flow conditions and the flow velocity, in particular with regard to a supersonic or subsonic flow, a variable flow channel cross-sectional shape in the entrance area of the diffuser can optimize the subsequent flow conditions in the jet engine and thus improve the flight behavior of the aircraft. A variable nozzle shape at the inlet of the diffuser can also enable higher flow conditions, i.e. higher Mach numbers. For example, the flow cross-section should initially be reduced in the case of supersonic flow in order to slow down the flow to subsonic, to increase the static pressure and in some cases also the fluid density. During this transition from supersonic to subsonic, several oblique compression shocks occur at the beginning and then a vertical compression shock at the end of this tapering nozzle shape. The taper angle must not become too large in order to prevent a flow separation at the boundary layers. The boundary layers could be influenced with further constructive measures. Subsequent to this vertical shock wave there is a subsonic flow, the nozzle shape must now expand in order to achieve a pressure increase.

In addition, the flow channel shape of the diffuser, in particular downstream of variable nozzle shapes at the diffuser inlet, can not be ideally straight, but rather curved or bent. The curvatures can be caused, for example, by a fuselage shape or wing shape of the aircraft, to which the flow channel with the subsequent jet engine is structurally adapted. Several curvatures, possibly with turning points in the curvature courses, can occur.

Furthermore, the flow profile and the direction of flow in the diffuser are influenced by different cross-sectional shapes at the diffuser inlet and diffuser outlet. The cross-sectional shape at the diffuser outlet is often circular, since the inlet cross-section on the jet engine is also usually circular. The cross-sectional shape at the diffuser inlet, on the other hand, can depend on an upstream air inlet. The air inlet can, for example, have a rectangular shape. This change in the cross-sectional shape within the diffuser or the flow channel influences the respective flow profile and the flow direction.

All these flow phenomena and structural details of the flow channel influence the direction and the flow profile of the fluid flowing through. At the same time, it is important and decisive which direction of flow is at the entrance to the jet engine. In particular, the direction of flow should ideally take place in the axial direction of a rotating or non-rotating inlet cone or inlet cone in order to achieve a stagnation point flow as precisely as possible at the tip of the cone. With this ideal stagnant flow, an optimal flow through the jet engine, i.e. the fan, the low-pressure and high-pressure compressor stages with a subsequent transition into the combustion chamber, can take place. With this ideal stagnant flow, for example, different boundary layer courses, possibly with the risk of flow separation, can be avoided and a through-flow can be achieved under the design conditions of the jet engine.

It is possible that a non-ideal stagnation point flow is already predetermined by the flow course, in particular by a flow channel cross-sectional course at the end of the flow channel, that is to say immediately before the entry into the jet engine. This can then mean that a rotating inlet cone, which is connected to a fan, cannot ideally flow against.

From the DE 10 2006 044 968 A1 an inlet cone for a jet engine is known. The inlet cone is arranged concentrically in the air inlet area of the jet engine on a rotatably mounted shaft.

From the DE 3026337 C2 a supersonic inlet for jet engines is known. The supersonic inlet has a tapering central body which is pivotably arranged on a trailing body arranged behind it.

From the EP 2 151 378 A2 a method for improving the flow conditions on the fan of an aircraft engine and a hub cone formed accordingly is known. To improve the flow conditions on the fan, at the stagnation point of the air at the tip of the hub cone, part of the incoming air flow is directed into the interior of the hub cone via air inlet openings and via air outlet openings in the area with the lowest static pressure at the downstream end of the hub cone The circumference is blown into the thick boundary layer on the hub cone in the direction of flow at a speed corresponding to the speed of the incoming air flow. The boundary layer is accelerated and its speed increased to that of the incoming air flow.

One object of the present invention is to propose a jet engine with an inlet cone in order to achieve the most ideal flow possible. Another object of the present invention is to propose a jet engine arrangement with a flow channel upstream of a jet engine.

The object according to the invention is achieved by a jet engine with the features of claim 1. Furthermore, the object according to the invention is achieved by a jet engine arrangement having the features of claim 5.

According to the invention, a jet engine, at least comprising an engine core area, a housing, an air inlet area and an inlet cone in the air inlet area of the jet engine is proposed. The jet engine can be designed as a turbofan engine. The inlet cone is connected to a non-rotatable component. A non-rotatable component means that it is a component that is not connected to a component that rotates during operation of the jet engine, for example with the blades of a fan, the blades of the low-pressure compressor, the high-pressure compressor, the low-pressure turbine or the high-pressure turbine. The entry cone can be connected to a guide grille. The guide grille can be connected to the housing at least in sections. The inlet cone is designed as an oblique inlet cone, the tip of the cone being arranged eccentrically to the axis of rotation of the jet engine.

A cone has a surface area and a base area. The straight line that runs through the center point of the base circle and the tip of the cone, i.e. the tip of the cone, is called the axis of the cone. In the case of an oblique cone, the axis is not perpendicular to the base. Therefore, the axis of the jet engine according to the invention does not run parallel to the axis of rotation of the jet engine. The axis of rotation can be referred to as the axis of rotation of the rotor of the jet engine. The rotor comprises at least one rotor stage, for example the low-pressure compressor or the high-pressure compressor.

The cone can be called a cone.

The cone can have straight and / or curved surfaces, at least in sections. The surfaces can be convex.

The guide grille can be or be referred to as an inlet grille.

An oblique entry cone can be referred to as a non-concentric entry cone. The outer surface of the inclined entry cone is not arranged concentrically around the axis of the entry cone. Thus the apex of the cone is not arranged concentrically either.

An oblique entry cone can be referred to as an asymmetrical entry cone or an asymmetrical entry cone. The tip of the asymmetrical inlet cone of the jet engine according to the invention is arranged outside the axis of rotation of the jet engine.

An oblique entry cone has non-rotationally symmetrical lateral surfaces. An oblique entry cone can be said to be asymmetrical.

The base of the inclined entry cone can be circular, elliptical or some other shape. An oblique entry cone with a circular base can be referred to as an oblique circular cone. An oblique elliptical cone can be described as a generalization of the oblique circular cone. An oblique elliptical cone with semiaxes of equal length of the elliptical base can be referred to as an oblique circular cone.

According to the invention, a jet engine arrangement is also proposed, comprising a jet engine according to the invention, an air inlet and a flow channel connecting the air inlet and the jet engine.

Advantageous further developments of the present invention are each the subject matter of subclaims and embodiments.

Exemplary embodiments according to the invention can have one or more of the features mentioned below in any combination, provided that one or the specific combination is not recognizable as obviously technically impossible for the person skilled in the art. The subjects of the subclaims also each specify exemplary embodiments according to the invention.

In all statements made above and below, the use of the expression “can be” or “may have” etc. is to be understood as synonymous with “is preferably” or “preferably” etc. and is intended to explain exemplary embodiments according to the invention.

Whenever alternatives with “and / or” are introduced herein, the person skilled in the art understands the “or” contained therein preferably as “either or” and preferably not as “and”.

Embodiments mentioned herein are to be understood as purely exemplary embodiments of the present invention according to the invention, which are not to be understood as restrictive.

A jet engine can be referred to as a turbine jet engine, turbo jet engine, turbo air jet engine, turbine air jet engine, gas turbine aircraft engine, jet engine or an aircraft engine. The jet engine comprises a gas turbine as a central component. Sometimes the jet engine is collectively referred to as a gas turbine.

A jet engine can be a single-flow jet engine or a turbofan engine, among others. The turbofan engine can be referred to as a bypass engine, a turbofan jet engine, a turbocharged turbine air jet engine or a fan engine. In a turbofan engine, an external air flow envelops an internal core flow through the gas turbine, in which the thermodynamic cycle takes place. Air is supplied and / or sucked in at the inlet of the turbofan engine.

In some embodiments of the invention, the axis of the entry cone is at an angle α arranged to the axis of rotation of the jet engine, the angle α can have a value between one degree and thirty degrees. The angle α can furthermore have a value between one degree and fifteen degrees, furthermore a value between one degree and ten degrees. An angle α a value of zero would correspond to a central flow, in this case the inlet cone would be straight. The choice of the angle α the structural design of the inlet cone depends in particular on the probable angle of attack of the jet engine. Since this angle of attack depends on many individual factors such as the respective operating state of the aircraft with the jet engine, i.e. during take-off, landing, continuous flight at a certain altitude, individual flight maneuvers or the speeds of the aircraft, for example subsonic or supersonic flight, it is often difficult to find this angle α to be precisely defined in advance. However, especially in aircraft that temporarily fly supersonic, there is a flow that may deviate from a central flow due to the upstream diffusers of the jet engine. This can be caused by a curved shape of the diffuser and / or certain flow phenomena within the diffuser. An approximate determination and definition of the angle α can be approximately determined with the help of experimental and / or computational methods or in another way. An angle α , which corresponds to the presumably actual direction of flow, can advantageously enable the jet engine to operate in an optimal manner in terms of flow. In this way, for example, fuel consumption can be reduced, the noise level can be reduced and / or the flight characteristics of the aircraft can be optimized.

In some embodiments according to the invention, the cone angle of the oblique entry cone has a value between twenty degrees and forty-five degrees, in particular between thirty degrees and forty degrees. The cone angle, both with an ideal flow parallel to the axis of the inlet cone and with a non-ideal flow, can have a considerable influence on the flow downstream of the inlet cone. At a low value, for example at a value of twenty degrees, at which the cone angle can be referred to as a flat cone angle, the flow deflection is less than with a larger cone angle. This can have a more favorable influence on the formation of the boundary layer and reduce the risk of a subsequent flow stall. On the other hand, the risk of solids entering the flow, for example stones, sand or birds, directly into the gas turbine, that is to say into the engine core area, can be greater with a shallow cone angle. A deflection of solids into the sheath flow area is often desirable, since the risk of damage to the low-pressure or high-pressure compressor, for example, is lower here. A deflection of solids into the sheath flow area may be more likely at an angle greater than approximately thirty-five degrees. The Flow conditions at the inlet cone and an influence on the flow downstream of the inlet cone essentially from the angle α between the axis of rotation of the jet engine and the axis of the oblique entry angle.

In some embodiments according to the invention, the entry cone is produced at least in sections by means of additive manufacturing. This can be advantageous because an oblique entry angle is generally not rotationally symmetrical and therefore manufacturing can be difficult. Furthermore, materials can be combined with one another in a simple manner by means of an additive manufacturing process, for example in order to use different materials in sections on the surface and / or at the tip of the entry cone. Likewise, the shape of the tip of the inlet cone, for example rounded for an ideal stagnation point flow, and / or the surface structure for a targeted influencing of the boundary layer flow, can be improved by means of greater design freedom in additive manufacturing.

In some embodiments according to the invention, the jet engine arrangement according to the invention has a flow channel in which a first cross section at the air inlet of the flow channel, the inlet cross section, is different from a second cross section at the inlet to the jet engine, the outlet cross section of the flow channel. In this embodiment, the air inlet is a separate component that is connected or arranged upstream of the inlet cross section of the flow channel. The air inlet can have certain functional and / or structural requirements. For example, it can functionally absorb the largest possible amount of air from the environment in order to increase the throughput and thus the thrust of the jet engine as much as possible. The air inlet can be formulated as a structural requirement for the simplest, most secure and adapted attachment to an aircraft fuselage. Depending on the expected operating conditions, for example in supersonic flight, further requirements can be placed on the air inlet. The outlet area of the flow channel is often adapted to the inlet cross section of the jet engine.

In some embodiments according to the invention, the first cross section of the flow channel, the inlet cross section, is rectangular or oval and the second cross section, the outlet channel, is circular. A rectangular inlet cross-section can easily be adapted to and connected to an air inlet which is likewise rectangular and is arranged upstream. In the case of a round, in particular circular exit cross-section, which may be necessary due to the shape of the downstream jet engine, the flow channel must transform or transfer the flow from a rectangular profile to a round profile with as little loss as possible. It should be noted that the flow conditions can be very different depending on the entry speed into the flow channel, in particular with regard to a supersonic flow and a subsonic flow.

In some embodiments according to the invention, the axis of the inlet cone essentially corresponds to the direction of flow of a fluid, the fluid generally being the ambient air flowing into the jet engine from the flow channel. In the case of an ideal flow with an identical flow direction of the fluid and the axis of the inlet cone, the central flow hits the tip of the inlet cone. In this case, the tip of the inlet cone corresponds to the stagnation point of the inlet cone. In this way, the inlet cone could ideally be flowed around, which means, for example, that the boundary layer flows at the inlet cone are homogeneous and the risk of flow separation can be minimized. However, it is often difficult to determine the exact direction of flow towards the inlet cone. This direction of flow towards the inlet cone essentially corresponds to the direction of flow at the outlet of the flow channel arranged upstream. As already discussed in detail above, the flow direction and the flow profile in the flow channel depend on numerous parameters and influencing factors.

In some embodiments according to the invention, the flow channel is curved at least in sections in its longitudinal direction. The curvatures can be single or multiple with points of inflection in the line of curvature. The curvatures can be two-dimensional in a plane or three-dimensional. Such curvatures can be due to the design, for example in order to adapt the flow channel as optimally as possible to an aircraft fuselage with the best possible flow. However, the curvatures should not lead to flow separation and / or turbulence within the flow channel. This is often difficult to predict in advance for the different operational situations of an aircraft. This can possibly be approximated through flow simulations on the computer and / or through experimental tests. Depending on the planned operational situations of an aircraft with supersonic and subsonic flights, possible predominant flow profiles and flow directions, in particular at the outlet cross-section of the flow channel, can be determined. As a result, the alignment of the axis of the inclined inlet cone on the jet engine could be established at least approximately advantageously.

In some embodiments according to the invention, the flow channel is designed as a diffuser for a supersonic flow and / or subsonic flow flowing in at the air inlet. In a diffuser, supersonic flows at the diffuser inlet can be slowed down to subsonic speeds in the diffuser with a simultaneous static pressure increase. Numerous changes in flow directions and flow conditions can occur in the diffuser.

In some embodiments according to the invention, the diffuser has movable wall segments that change the flow cross-section, in particular in or near the inlet region of the diffuser. Depending on the flow conditions and the flow velocity, in particular with regard to a supersonic or subsonic flow, a variable flow channel cross-sectional shape by means of variable wall segments in the entrance area of the diffuser can optimize the subsequent flow conditions in the jet engine and thus improve the flight behavior of the aircraft. A variable nozzle shape at the inlet of the diffuser can enable higher flow conditions, i.e. higher Mach numbers. For example, the flow cross-section should initially be reduced in the case of a supersonic flow in order to slow down the flow to subsonic, to increase the static pressure and in some cases also the fluid density. During this transition from supersonic to subsonic, several oblique compression shocks occur at the beginning and then a vertical compression shock at the end of this tapering nozzle shape. The taper angle must not become too large in order to prevent a flow separation at the boundary layers. The boundary layers could be influenced with further constructive measures. Subsequent to this vertical shock wave, there is usually a predominantly subsonic flow; the nozzle shape must now expand in order to achieve a pressure increase.

In some embodiments according to the invention, the movable wall segments enable a cross-section reduction at the diffuser inlet for a supersonic flow with a downstream cross-sectional enlargement for a subsonic flow.

In some embodiments according to the invention, the flow channel has at least one guide grille. A guide grille can advantageously generate a desired flow profile and / or a desired flow direction. Flow losses and / or turbulence and / or flow separation with subsequent instabilities in the flow should be avoided.

In some embodiments according to the invention, a guide grille at the outlet of the flow channel or of the diffuser is designed for a flow direction parallel to the axis of an inclined inlet cone arranged downstream in the jet engine.

• 1 zeigt ein erfindungsgemäßes Strahltriebwerk als Mantelstromtriebwerk in einer Schnittebene;
• 2a zeigt exemplarisch einen schiefen Eintrittskegel eines erfindungsgemäßen Strahltriebwerks;
• 2b zeigt exemplarisch einen weiteren schiefen Eintrittskegel mit konvex gekrümmten Oberflächen eines erfindungsgemäßen Strahltriebwerks; und
• 3 zeigt eine erfindungsgemäße Strahltriebwerksanordnung mit einem erfindungsgemäßen Strahltriebwerk, einem Diffusor als Strömungskanal und einem Lufteinlass.

The present invention is explained below by way of example with reference to the accompanying drawings, in which identical reference symbols designate identical or similar components. In the highly schematically simplified figures, the following applies:

• 1 shows a jet engine according to the invention as a turbofan engine in a sectional plane;
• 2a shows an example of an inclined inlet cone of a jet engine according to the invention;
• 2 B shows an example of a further inclined inlet cone with convexly curved surfaces of a jet engine according to the invention; and
• 3 shows a jet engine arrangement according to the invention with a jet engine according to the invention, a diffuser as a flow channel and an air inlet.

1 shows a jet engine according to the invention 100 as a turbofan engine in a sectional plane. The turbofan engine is characterized in that it is downstream of a fan 1 near the air inlet area 3 the airflow is separated. An outside stream of air 5 , the so-called bypass flow, is applied to an engine core area 9 passed by and between a housing 7th and the engine core area 9 to the outlet area 11 of the jet engine 100 guided. An inner core stream 13th is usually by a low pressure compressor, a high pressure compressor, a combustion chamber, a high pressure turbine and a low pressure turbine, which in 1 are not shown in detail, to the outlet area 11 of the jet engine 100 guided. The ratio between the outside airflow 5 , which is also called bypass flow or bypass flow, to the inner core flow 13th , varies depending on the embodiment of the jet engine 100 between 0.3 and 14, with the lower value being used, for example, for supersonic aircraft and the higher value for wide-body aircraft in the subsonic range. The inner core stream 13th provides the so-called wave power, which in turn, among other things, the fan 1 drives. Furthermore, in 1 a suspension 15th of the jet engine 100 shown on a wing or on an aircraft fuselage.

Furthermore, according to the invention, it is upstream of the fan 1 an oblique entry cone 17th shown, which has a guide grille 19th firmly to the housing 7th connected is. The fan 1 as well as the impellers of the compressor and the turbine rotate around an axis of rotation 23 , which can be referred to as the axis of rotation of the jet engine. Rotationally symmetrical, rotatable entry cones which rotate together with a rotatable fan are known from the prior art. An inclined flow towards the jet engine 100 can have different causes, for example an inclined arrangement of the jet engine 100 on the plane. In particular, there can be an inclined flow 21 occur in an upstream flow channel, as shown in 3 is shown and described in more detail.

The geometry of the angled entry cone 17th is in the 2a and 2 B described in more detail.

2a shows an example of an inclined entry cone 17th of a jet engine according to the invention 100 . The entrance cone 17th can, as in 1 shown, have a paragraph on which optionally a guide grille 19th or struts for attaching to the housing 7th are attached.

The crooked entrance cone 17th has an asymmetrical shape with respect to the axis of rotation 23 on. The asymmetry is due to an angle other than zero α between the axis of rotation 23 and the axis 27 of the entry cone 17th marked. The axis 27 of the entry cone 17th runs through its tip by definition 29 and the center point 31 the base area 33 of the entry cone 17th . The base 33 can take any shape, especially circular.

The cone angle 25th , which is the outer surface 35 spans, can be referred to as the opening angle.

In the case of a symmetrical entry cone, the two axes would be the axis of rotation 23 and the axis 27 of the entry cone 17th coincide. This would correspond to a straight entry cone.

With a flow in the direction of the axis 27 of the entry cone 17th can the flow to the tip 29 are referred to as stagnation point flow.

The exemplary embodiment in 2a has a straight lateral surface 35 on.

2 B shows an example of another inclined entry cone 17th with a convex curved surface 35 of a jet engine according to the invention 100 .

Further exemplary lateral surfaces 35 may be concave or other shape.

3 shows a jet engine arrangement according to the invention 200 with a jet engine according to the invention 100 , a diffuser as a flow channel 300 and an air inlet 400 . The supplied and / or sucked in air mass flow 37 first enters the air inlet 400 one, is then passed through the flow channel 300 guided and then as a flow 21 the jet engine 100 forwarded.

Jet engines 100 with an upstream diffuser as a flow channel 300 are used in particular in aircraft that sometimes fly supersonic, i.e. with a Mach number greater than one. The Mach number can be defined as the ratio of the speed of the aircraft to the speed of sound of the surrounding fluid.

The geometric and structural design of the diffuser 300 can depend on and be influenced by many factors. For example, the diffuser 300 perpendicular to the display plane in 3 Curved or double-curved in an S-shape so that, for example, an adaptation to the aircraft fuselage of a supersonic jet is possible. Furthermore, it can be displayed in 3 , as shown, be curved in order to achieve the highest possible air mass inflow 38.

As already discussed in detail above, an entry of an air mass inflow 38 at supersonic speed must occur before an oncoming flow 21 into the jet engine 100 decelerated to subsonic. This is usually done initially by a cross-sectional tapering in the entry area of the air inlet 400 . Movable wall segments can be used for this 39 , for example designed as flaps, or the like can be provided structurally at the diffuser inlet. However, this tapering of the cross-section only makes sense in the case of supersonic, not subsonic, in order to increase the static pressure in the flow. In the case of a supersonic flow, oblique shock waves first occur with a straight shock shock at the end at the narrowest cross-section. After this straight shock wave, the air then usually continues to flow in subsonic form. Depending on the local flow direction, possible flow detachments on the movable flaps, the respective boundary layer courses or other influencing factors, the direction of the flow can be influenced and it may not be ideally guided in the longitudinal direction in the diffuser. This can lead to the inflow into the jet engine, particularly if the diffuser end area is not aligned straight 100 not in the direction of rotation of the rotor in the jet engine 100 he follows. In this case it would be the entrance cone in the jet engine 100 be straight, the flow could probably not be evenly distributed over the cross-section. This can in particular lead to an undesired inflow into the inner core flow 13th lead, with subsequent detachments, boundary layer influences and / or turbulence in the inner core flow 13th . This would at least lead to a reduced performance of the jet engine. A jet engine according to the invention 100 with an inclined entry cone 17th can prevent this advantageously. The determination of an approximately optimal alignment of the tip 29 of the entry cone 17th and thus the stagnation point of the flow 21 can be at least approximately determined in various ways, as discussed above. These include, for example, flow simulations and / or experimental tests based on the actual structural design of the diffuser.

The diffuser can, in particular in the area downstream of the movable wall segments, have one or more guide elements for aligning the flow.

List of reference symbols

a

axial; Axial direction

r

radial; Radial direction

α

Angle between the axis of rotation and the axis of the entry cone

100

Jet engine

200

Jet engine arrangement

300

Flow channel; Diffuser

400

Air inlet

1

fan

3

Air inlet area

5

external airflow; Sheath current

7th

casing

9

Engine core area

11

Outlet area

13th

inner core flow

15th

Suspension of the jet engine

17th

inclined entry cone

19th

Baffle

21

Inflow

23

Axis of rotation

25th

Cone angle

27

Axis of the entry cone

29

Tip of the entry cone

31

Center of the base of the entry cone

33

Base area of the entry cone

35

Lateral surface of the entry cone

37

Air mass flow

39

movable wall segments at the diffuser inlet

QUOTES INCLUDED IN THE DESCRIPTION

This list of the documents listed by the applicant was generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Patent literature cited

• DE 102006044968 A1 [0007]
• DE 3026337 C2 [0008]
• EP 2151378 A2 [0009]

Claims (14)

Jet engine (100), in particular designed as a turbofan engine, comprising an air inlet area (3), a housing (7), an engine core area (9) and an inlet cone (17) in the air inlet area (3), the inlet cone (17) having a non-rotatable Component, in particular connected to a guide grille (19), characterized in that the inlet cone (17) is designed as an oblique inlet cone (17), the cone tip (25) being arranged eccentrically to the axis of rotation (23) of the jet engine (100). Jet engine (100) after Claim 1 , the axis (27) of the inlet cone (17) being arranged at an angle α to the axis of rotation (23) of the jet engine (100), the angle α having a value between one degree and thirty degrees, in particular a value between one degree and fifteen Degree, further in particular a value between one degree and ten degrees. Jet engine (100) after Claim 1 or 2 wherein the cone angle (25) of the entry cone (17) has a value between twenty degrees and forty-five degrees, in particular between thirty degrees and forty degrees. Jet engine (100) according to one of the preceding claims, wherein the inlet cone (17) is manufactured at least in sections by means of additive manufacturing. A jet engine arrangement (200) comprising a jet engine (100) according to one of the Claims 1 until 4th , an air inlet (400) and a flow channel (300) connecting the air inlet (400) and the jet engine (100). Jet engine assembly (200) according to Claim 5 , wherein a first cross section of the flow channel (300) on the air inlet (400) is different from a second cross section of the flow channel (300) on the jet engine (100). Jet engine assembly (200) according to Claim 6 , wherein the first cross section is rectangular or oval and the second cross section is circular. Jet engine arrangement (200) according to one of the Claims 5 until 7th , the axis (27) of the inlet cone (17) essentially corresponding to the direction of flow of a fluid which flows from the flow channel (300) into the jet engine (100). Jet engine arrangement (200) according to one of the Claims 5 until 8th , wherein the flow channel (300) is curved at least in sections in the longitudinal direction. Jet engine arrangement (200) according to one of the Claims 5 until 9 , wherein the flow channel (300) is designed as a diffuser for a supersonic flow and / or subsonic flow flowing in at the air inlet (400). Jet engine assembly (200) according to Claim 10 , wherein the diffuser has movable wall segments (39) which change the flow cross section. Jet engine assembly (200) according to Claim 11 , wherein the movable wall segments (39) enable a cross-section reduction at the diffuser inlet for a supersonic flow with a downstream subsequent increase in cross-section for a subsonic flow. Jet engine arrangement (200) according to one of the Claims 5 until 12th , wherein the flow channel (300) has at least one guide element. Jet engine assembly (200) according to Claim 13 , wherein the guide element is designed at its outlet for a flow direction parallel to the axis (27) of the inlet cone (17).

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