DE102019210417B4 - Missile and method for operating a missile - Google Patents

Abstract

Missile (1), with a wing (2), seen in section, on the one hand by a lower first profile surface (4) when the missile (1) is operated as intended and on the other hand by an upper one at a wing transition point (6) with the first profile surface ( 4) converging second profile surface (5), the first profile surface (4) enclosing at least one air inlet opening (7) and the second profile surface (5) enclosing at least one air outlet opening (8), and the missile (1) has a propulsion device (12). an air conveying device (1) which is provided and designed for sucking in air through the at least one air inlet opening (7) and for discharging the sucked air through the at least one air outlet opening (8), the at least one air outlet opening (8) being at least partially is overlapped by a deflection element (15) which is in flow connection with the second profile surface (5) with the air outlet opening (8). end air outlet gap (16), and the air outlet opening (8) is fluidically connected to the air outlet gap (16) via a connecting channel (17) which, seen in section, is present between the second profile surface (5) and the deflection element (15), the second profile surface (5) has a first area starting from the wing transition point (6) and a second area adjoining the first area and delimiting the air outlet gap (16), wherein - viewed in section - the first area is curved and the second area is curved or is planar, and wherein the connecting duct (17) has a cross section that decreases in the direction of the air outlet gap (16), so that it is designed like a nozzle, wherein the shape of the connecting duct (17) is selected in such a way that the air in the Air outlet gap (16) has a flow rate lying in the subsonic range, characterized in that the fewest s an air inlet opening (7) and the at least one air outlet opening (8) are fluidically connected via a flow channel (9) formed centrally in the wing (2) in section, with a ducted propeller (10) of the air conveying device (11 ) is arranged to be rotatable about an axis of rotation.

Description

The invention relates to a missile according to the preamble of claim 1 and a method for operating a missile according to the preamble of claim 10.

From the prior art, for example, the publication DE 20 2018 104 722 U1 known. This describes an aircraft with a frame structure and a plurality of lifting rotors arranged on the frame structure, by means of which a primary lift directed vertically upwards and propulsion forces can be generated. In this case, it is provided that a jet turbine is also provided, the thrust jet of which can be aligned in such a way that a secondary lift which is directed essentially parallel to the primary lift can be generated and which can be superimposed on the primary lift.

Furthermore, the reference discloses DE 20 31 281 A An apparatus for generating vertical aerodynamic lift forces, comprising a levitation body having a substantially frusto-conical upper surface formed thereon, a chamber arrangement cooperating with the upper surface to form a biconical annular nozzle, and having a pressurized medium at supersonic velocity through the Device discharging nozzle over the upper surface area of the float.

Document WO 2009/068835 A1 shows a static wing for an aircraft, comprising: a ring portion with a wing-shaped cross section, wherein a leading edge of the wing defines an inner circumference of the ring and a trailing edge of the wing defines the outer circumference of the ring; and a diffuser plate provided over the opening of the ring portion, the diffuser plate being larger than the opening of the ring portion and having an upper surface whose curvature substantially follows the curvature of the annular portion and a lower surface which is similarly curved.

Finally describes the U.S. 4,674,708 A an aircraft having a disc-shaped fuselage having a planar top surface and including at least one vertical duct extending longitudinally therethrough, an impeller disposed in the duct for generating an upward flow of gas, and a gas flow straightener positioned relative to the Channel is movably mounted. The vertical channel is located in the center of the fuselage. The impeller is a vertical turbo engine. The gas regulator includes a central conical member, an upper annular flow control member, and a plurality of radially spaced outflow control members. The outflow control elements are hydraulic cylinders. A flap is fitted around the bottom opening of the channel. The flap is movable so that in one position it seals the channel against the ingress of liquids. An air passage extends into and communicates with the duct.

It is the object of the invention to propose a missile which has advantages over known missiles, in particular can be scaled as desired with regard to its payload and also enables particularly efficient operation.

According to the invention, this is achieved with a missile having the features of claim 1 . This has a wing which, seen in section, is delimited on the one hand by a first profile surface which is lower when the missile is operated as intended and on the other hand by an upper second profile surface which converges with the first profile surface at a wing surface transition point, the first profile surface having at least one air inlet opening and the second Profile surface encloses at least one air outlet opening and the missile has a drive device with an air conveying device that is provided and designed for sucking in air through the at least one air inlet opening and for discharging the sucked air through the at least one air outlet opening, wherein the at least one air outlet opening is at least partially covered by one Deflection element is overlapped, which limits a standing with the air outlet in flow communication air outlet gap with the second profile surface, and the air outlet with the air outlet gap is fluidically connected via a connecting channel, which is present in section between the second profile surface and the deflection element.

Furthermore, the second profile surface has a first area starting from the wing transition point and a second area adjoining the first area and delimiting the air outlet gap, wherein - viewed in section - the first area is curved and the second area is curved or flat, and wherein the connecting duct has a cross section that decreases in the direction of the air outlet gap, so that it is designed like a nozzle, the shape of the connecting duct being selected such that the air in the air outlet gap has a flow velocity in the subsonic range.

According to the invention, the at least one air inlet opening and the at least one air outlet opening have a central cross-section flow channel formed on the airfoil, with a ducted propeller of the air conveying device being arranged in the flow channel so that it can rotate about an axis of rotation.

In principle, the missile can have any configuration, for example it is in the form of an unmanned aerial vehicle or—preferably—an aircraft. The aircraft means a vehicle that flies within the earth's atmosphere. In this respect, it represents a mobile means of transport that is used to transport people, goods or the like. The aircraft is therefore intended and designed for passenger transport and/or freight transport. The missile or the aircraft is heavier than air and has the propulsion device or a power drive, so that the missile as a whole can be referred to as an aircraft. Generally speaking, an airplane is a heavier-than-air aircraft that uses non-rotating lifting surfaces to generate the dynamic lift necessary to fly it.

The flying body according to the invention has the airfoil, in which case, of course, a plurality of airfoils can also be used, which are preferably connected to one another, in particular via a superstructure of the flying body. In the latter case, the following explanations for the wing preferably apply to several of the wing or to each of the wing. Viewed in section, the wing is delimited by two profile surfaces, namely the first profile surface and the second profile surface. To this extent, the first profile surface and the second profile surface delimit the wing in opposite directions, in particular with respect to a longitudinal central axis of the wing or of the missile. When the missile is operated as intended, the first profile surface is arranged at the bottom and the second profile surface is arranged at the top, the first profile surface is therefore on the side of the missile facing a ground, in particular a ground, whereas the second profile surface is on a side of the missile facing away from the ground is arranged.

Viewed in section, the first profile surface and the second profile surface converge at the airfoil interface and therefore merge into one another at the airfoil interface. The wing transition point is, for example, a line, in particular a line with a continuous course and/or a continuous line, or runs continuously along the line. The continuous line means that the line is self-contained, ie a starting point of the line corresponds to an end point of the line and the starting point and the end point are continuously connected to one another. For example, the line can be a straight line. However, it is preferably round or circular or oval. The airfoil interface may correspond to or be present on an airfoil outboard. The outer surface of the wing represents or is present at that point on the wing which is furthest away from the longitudinal central axis. For example, the wing transition point is a wing transition edge at which the first profile surface and the second profile surface meet discontinuously and merge into one another, seen in section.

While the first profile surface and the second profile surface converge at the wing transition point, they are spaced apart at a wing inner side opposite the wing transition point in section. The first profile surface and the second profile surface therefore run away from one another, viewed in section, starting from the wing transition point in the direction of the inside of the wing, so that they are at a greater distance from one another on the inside of the wing than at the point of transition of the wing. Viewed in section, the wing transition point is in any case further to the outside than the inside of the wing, in particular it is further spaced from the longitudinal central axis. Provision is particularly preferably made for the distance between the first profile surface and the second profile surface to increase steadily starting from the wing transition point in the direction of the inside of the wing. For this purpose, the two profile surfaces are configured and/or arranged accordingly.

The missile has the air inlet opening and the air outlet opening. During operation of the missile, air is sucked in through the air inlet opening and discharged through the air outlet opening. For this purpose, the air inlet opening and the air outlet opening are fluidically connected to one another, so that the air sucked in through the air inlet opening is subsequently discharged through the air outlet opening. It can be provided that there is only a single air inlet opening. Alternatively, several air inlet openings are implemented. This applies correspondingly to the air outlet opening, so that there is either a single air outlet opening or a plurality of air outlet openings. If the air inlet opening is discussed in the context of this description, the corresponding statements always apply to the at least one air inlet opening and vice versa and—in the case of several air inlet openings—to each of the air inlet openings. Analogously, the explanations for the air outlet openings always apply to the at least one air outlet opening and vice versa and—in the case of several air outlet openings—to each of the several air outlet openings. In this sense, the Formu lations air inlet opening and at least one air inlet opening on the one hand and air outlet opening and at least one air outlet opening on the other hand are each provided with the same meaning.

The air inlet opening is bordered by the first profile surface and the air outlet opening by the second profile surface. This means in particular that the air inlet opening is delimited by the first profile surface on the inside of the wing and the air outlet opening is delimited by the second profile surface on the inside of the wing. In particular, the first profile surface surrounds the air inlet opening and the second profile surface surrounds the air outlet opening in each case in a ring. However, it can also be provided that the air inlet opening extends through the first profile surface and/or the air outlet opening extends through the second profile surface.

The missile has the propulsion device, which in turn has the air conveying device. This is provided and designed to convey air from the inlet opening in the direction of the air outlet opening during the intended flight operation of the missile and to this extent to draw in air through the air inlet opening and expel it through the air outlet opening. The missile is now arranged geodetically in such a way that during normal flight operation of the missile the first profile surface is arranged at the bottom and the second profile surface is arranged at the top. In other words, the first profile surface faces the ground or the ground, whereas the second profile surface faces away from the ground or the ground. This means that the propulsion device conveys air from the underside of the missile to its upper side. The drive device is provided and designed to convey air from a first side of the wing, which is the underside, on the side of the first profile surface, to a second side of the wing, which is the upper side, on the side of the second profile surface.

In order to nevertheless generate lift and enable flight operations, the air outlet opening is at least partially overlapped by the deflection element. The deflection element serves to deflect the air exiting through the air outlet opening and, for this purpose, together with the second profile surface, delimits the air outlet gap, which is in flow connection with the air outlet opening. Since the deflection element delimits the air outlet gap together with the second profile surface, the air exiting the air outlet opening and then being fed to the air outlet gap exits through the air outlet gap at least partially parallel or completely parallel to the second profile surface. As a result, an air film is formed on the second profile surface, which film flows over the second profile surface in the direction of the wing transition point. The air outlet gap is therefore provided and designed in particular for the discharge of air parallel to the second profile surface and for the formation of an air film on the second profile surface.

At the latest at the wing transition point, the air film tears away from the second profile surface or from the entire missile, in order to then continue to flow in the form of a free jet air flow, namely in the direction of the ground. This free-jet air flow induces a vortex. On the inside of the wing, the intake of air through the air inlet opening supports the formation of the lifting vortex, in that the air is sucked into the air inlet opening in the form of an intake air flow. Viewed in section, the free jet air flow now flows on the one hand of the support vortex in a free jet direction and on the other hand of the support vortex the intake air flow in an intake direction, the free jet direction and the intake direction each being tangential to the support vortex and pointing in different directions, in particular opposite to one another. The support vortex is thus formed at least by the free jet air flow. In addition, the intake air flow can (optionally) contribute to the formation of the lifting vortex. The support swivel is preferably at least partially below the first profile surface, ie in particular between the first profile surface and the ground or the ground.

The tearing of the air film from the second profile surface takes place at a tear-off point, seen in section. This can correspond to the wing transition point or be spaced from it. On average, based on a distance between the inside of the wing or the air outlet gap and the wing transition point along the second profile area, the distance between the inside of the wing or the air outlet gap on the one hand and the tear-off point on the other hand, again along the second profile area, is at least 50%, at least 60%, at least 70% , at least 80% or at least 90%. The air film particularly preferably flows up to the wing transition point along the second profile surface.

The lift acting on the missile is achieved by different mechanisms of action. On the one hand, due to the discharge of the air from the air outlet gap, the air film is present on the upper side of the wing, which due to its high flow speed according to Bernoulli's equation causes a negative pressure compared to the underside of the wing. Namely because the flow speed of the air on the side of the second profile surface is greater than the flow speed of the air on the side of the first profile surface, the pressure present on the side of the second profile surface is lower than the pressure present on the side of the first profile surface. The pressure difference between the pressures on the first profile surface and on the second profile surface, ie the pressure difference between the underside and the upper side of the wing, already causes part of the lift acting on the missile. Another part of the lift is provided by the support vortex as soon as it is present. It is clear that the lifting vortex can only be formed when there is a sufficient distance between the missile and the ground. In order to use the lift generated by the support vortex, it is therefore first necessary to space the missile from the ground. This is preferably done using a mechanical lifting device which lifts the missile relative to the ground.

The use of the support vortex to at least partially provide the lift enables a particularly energy-efficient operation of the missile, because the generation and maintenance of the support vortex is possible with a comparatively small amount of energy, which is significantly lower than the energy expenditure that would be necessary to directly raise the missile. The gyroscopic vortex also results in a particularly high stability of the missile in the air, because the gyroscopic vortex occupies a large volume of air or creates a voluminous air cushion for the missile. The wing and to that extent the missile are arbitrarily scalable in terms of carrying capacity, because the support swivel is basically arbitrarily scalable. This is because the support vortex is formed in the manner of a potential vortex.

The support surface is preferably symmetrical when viewed in section, for example even rotationally symmetrical. In principle, two different configurations of the wing can be distinguished. According to a first embodiment, the wing is straight and, viewed in section, is configured symmetrically with respect to an axis of symmetry. Within the scope of this description, the axis of symmetry is also referred to as the longitudinal central axis. In section, the axis of symmetry preferably runs centrally through the air inlet opening and the air outlet opening. The first profile surface and the second profile surface thus extend, starting from the wing transition point in the direction of the axis of symmetry, up to the inside of the wing, which preferably delimits the air inlet opening and/or the air outlet opening. On the opposite side of the axis of symmetry, the first profile surface and the second profile surface in turn extend from a (further) wing inside to a (further) wing transition point, at which they in turn converge. To this extent, the wing transition points run at a distance from one another, in particular at a distance from one another in parallel. They preferably each run straight. Viewed in section, the missile is symmetrical with respect to the air inlet opening and/or the air outlet opening or with respect to the axis of symmetry running centrally through the air inlet opening and/or the air outlet opening. In such an embodiment, the support swivel is cylindrical.

However, a second configuration of the missile is preferred, according to which the wing is round or oval, so that the missile is designed, for example, in the manner of a flying disc. The round wing is preferably rotationally symmetrical with respect to the longitudinal center axis, with the longitudinal center axis representing a mirror axis of the wing when viewed in section. The wing transition point is preferably formed continuously in the circumferential direction with respect to the longitudinal central axis. The wing transition point preferably has a continuously constant curvature in the circumferential direction, so that the wing transition point ultimately runs in a circle. To this extent, the wing is ring-shaped with respect to the longitudinal central axis. In the case of the second embodiment of the missile, the support vortex is toroidal, with the torus being understood as the body of revolution of a circle, in particular as a body of revolution around the longitudinal central axis of the wing. Such a configuration of the support vortex that is closed in the circumferential direction enables a particularly energy-efficient flight operation of the missile.

A development of the invention provides that the at least one air inlet opening is arranged centrally in the first profile surface and/or the at least one air outlet opening is arranged centrally in the second profile surface. In other words, the air inlet opening or the air outlet opening are formed centrally in the respective profile surface in section, in particular in longitudinal section with respect to the longitudinal center axis. This means that the air flowing out of the air outlet opening when the missile is in flight flows over a large part of the second profile surface or even the entire profile surface, in particular in the form of the air film. For example, viewed on average, the air flows over at least 50%, at least 60%, at least 70%, at least 80% or at least 90% of the second profile surface. The mass flow of the Air film during its flow over the second profile surface. This means that the free-jet air flow, which ultimately generates the support vortex, at least temporarily has a larger mass flow than the air film directly when it emerges from the air outlet opening or the air outlet gap. This further improves the energy efficiency of the missile.

A development of the invention provides that the supporting surface is ring-shaped with respect to a longitudinal central axis and the profile surfaces are spaced apart from one another in axial direction in some areas, as seen in longitudinal section. This applies in particular to the second embodiment of the missile already mentioned above. According to this, the bearing surface is ring-shaped and preferably rotationally symmetrical with respect to the longitudinal central axis. Seen in longitudinal section with respect to the longitudinal central axis, the profile surfaces are spaced apart from one another in areas in the axial direction, namely away from the wing transition point at which they converge. With this configuration of the missile, the particularly energy-efficient flight operation already mentioned is made possible.

The invention provides that the at least one air inlet opening and the at least one air outlet opening are fluidically connected via a flow channel formed centrally in the wing in section, with a ducted propeller of the air conveying device being arranged in the flow channel so that it can rotate about an axis of rotation. The section is in turn preferably to be understood as the longitudinal section with respect to the longitudinal central axis of the wing. The flow duct connecting the air inlet opening and the air outlet opening to one another in terms of flow preferably has a constant flow cross section or a constant flow cross section area throughout. Ultimately, this means that the air inlet opening and the air outlet opening have the same flow cross-sectional area.

The air inlet opening is delimited continuously by the first profile surface in the circumferential direction with respect to the longitudinal central axis, so that the first profile surface forms an edge of the air inlet opening on the inside of the wing. Analogous to this, the second profile surface continuously encompasses the air outlet opening in the circumferential direction, so that the second profile surface forms an edge of the air outlet opening on the inside of the wing. The edge of the air inlet opening preferably lies continuously in an imaginary first plane, whereas the edge of the air outlet opening preferably lies continuously in an imaginary second plane. The two imaginary planes are in particular arranged parallel to one another and are preferably each perpendicular to the longitudinal center axis. Correspondingly, a normal direction of the air inlet opening and a normal direction of the air outlet opening are arranged parallel to the longitudinal center axis or coincide with it.

The ducted propeller of the air conveying device is arranged and rotatably mounted in the flow channel. The ducted propeller is provided and designed to suck in the air through the air inlet opening during flight operation of the missile and to convey it in the direction of the air outlet opening, so that the air sucked in through the air inlet opening subsequently exits from the air outlet opening. The ducted propeller represents an impeller of the air conveying device, which is designed as an axial impeller. In addition to the ducted propeller, the air conveying device can have a guide wheel which is arranged downstream of the ducted propeller and serves to align the air flowing through it parallel to the longitudinal center axis. The ducted propeller is arranged overall in such a way that on the one hand it works extremely efficiently because flow losses at blade tips of the ducted propeller are reduced and on the other hand it is reliably protected against damage.

According to the invention, it is provided that the air outlet opening is fluidically connected to the air outlet gap via a connecting channel which, seen in section, is present between the second profile surface and the deflection element. The connecting channel is delimited in section, in particular in longitudinal section with respect to the longitudinal center axis, on the one hand by the second profile surface and on the other hand by the deflection element, ie on opposite sides. The connecting duct preferably runs at an angle with respect to the aforementioned flow duct, so that a main flow direction of the air flowing through the connecting duct is angled with respect to a main flow direction of the air flowing through the flow duct, i.e. enclosing an angle that is greater than 0° and less than 180°. The angle is preferably at least 45° and at most 135°, at least 60° and at most 120°, at least 70° and at most 110°, at least 80° and at most 100° or approximately or exactly 90°.

Like the air outlet gap, the connecting channel is preferably continuous and in particular uninterrupted in the circumferential direction with respect to the longitudinal center axis, so that the air film on the second profile surface is also continuous and uninterrupted. As a result, the support vortex is generated particularly efficiently. Of course, however, thin webs can be present, by means of which the deflection element connected to the wing. From a fluidic point of view, these webs are designed in such a way that it can still be assumed that there is an uninterrupted connection channel and an uninterrupted air outlet gap.

According to the invention, the connecting duct has a cross section that increases or decreases in the direction of the air outlet gap, so that it is designed in the manner of a nozzle. Depending on the flow speed of the air in the flow channel, the connecting channel can either widen or narrow in the direction of the air outlet gap. The shape of the connecting duct is selected in particular in such a way that the air in the air outlet gap or the air film has a desired flow rate immediately after it emerges through the air outlet gap. According to the invention, this flow speed is in the subsonic range, so that no negative mechanical influences on the missile due to fluidic shocks or the like are to be expected. The configuration of the connecting channel in the manner of a nozzle enables effective operation of the missile.

A development of the invention provides that the air outlet gap is designed to be continuous, in particular in the circumferential direction with respect to the longitudinal center axis. This has already been pointed out above. Such a configuration enables the air film to be produced uniformly and accordingly enables the support vortex to be excited in a particularly efficient manner.

The invention also provides that the second profile surface has a first area starting from the wing transition point and a second area adjoining the first area and delimiting the air outlet gap, wherein—seen in section—the first area is curved, in particular continuous, and the second area is curved or planar. The second profile surface thus has the first area and the second area, in particular it is composed exclusively of the first area and the second area. Preferably, the first area and the second area are each ring-shaped, with the first area encompassing the second area in the circumferential direction with respect to the longitudinal center axis. The first area immediately follows the second area, in particular the first area transitions seamlessly and steadily into the second area.

Viewed in section, in particular in longitudinal section with respect to the longitudinal center axis, the first region of the second profile surface is curved, specifically in the radial direction with respect to the longitudinal center axis outwards away from the deflection element. This means that the distance in the axial direction between the second profile surface and the deflection element in the first region also increases as the distance in the radial direction increases. Particularly preferably, the first region is continuously curved, that is to say has a non-zero curvature throughout, as viewed in section or longitudinal section.

The second area can also be curved, for example partially or continuously. However, it can also be planar throughout or--alternatively--partially curved and partially planar. For example, seen in section, the second region is in overlap with the deflection element. On the other hand, the first region is located to the side or outside of the deflection element in the radial direction. The division of the second profile surface into the first area and the second area enables a particularly advantageous air flow, so that the air film sweeping over the second profile surface generates the lifting vortex in a particularly effective manner after it is torn off the wing.

A further development of the invention provides that the first profile surface, seen in section, is set back in some areas in relation to the wing transition point in the direction of the second profile surface, so that a turbulence chamber is formed, which is encompassed by the wing transition point. The section is in turn preferably to be understood as the longitudinal section with respect to the longitudinal center axis. The partial setting back of the first profile surface with respect to the wing transition point is realized, for example, by a curvature of the first profile surface and/or a recess, ie in particular the formation of a step. In the case of the curvature, a radius of curvature of the curvature is preferably chosen in such a way that it is larger than a radius of the generated lifting vortex or corresponds approximately or exactly to it.

The vortex chamber is bounded axially upwards by the first profile surface and radially outwards by the airfoil interface. The turbulence chamber preferably extends inward in the radial direction as far as the air inlet opening. The vortex chamber serves to receive the gyroscopic vortex in certain areas during flight operation of the missile and effects stabilization and positioning of the gyroscopic vortex, particularly in the radial direction. Due to the positioning of the support vortex with the aid of the vortex chamber, a particularly effective generation of lift is achieved.

A further development of the invention provides that the first profile surface and the second profile surface are curved in the same direction, at least in certain areas, when viewed in section. Curvature in the same direction is to be understood as meaning a curvature with the same sign. In this way, a particularly material- and weight-saving configuration of the missile is achieved. For example, it is provided that the first profile surface has a continuously constant curvature, starting from the wing transition point up to the inside of the wing or up to the air inlet opening. Additionally or alternatively, the first profile surface has a curvature that changes, in particular decreases, starting from the wing transition point in the direction of the inside of the wing.

For example, the curvature of the first profile surface directly at the wing transition point is greater than the curvature of the second profile surface. In the direction of the inside of the wing, on the other hand, the curvature of the first profile surface decreases, for example down to a curvature of 0. Particularly preferably, the first profile surface, viewed in section or longitudinal section, has a straight course at a distance from the wing transition point, with this straight course of the first profile surface receiving imaginary plane is preferably perpendicular to the longitudinal center axis. In the area of the straight run, the first profile surface particularly preferably delimits the air outlet gap. This achieves a particularly effective air flow.

A further development of the invention provides that the second profile surface merges into the first profile surface at the wing transition point at an angle which is at least 0° and at most 60° with respect to a straight line perpendicular to an imaginary plane continuously accommodating the wing transition point. This applies in particular when viewed in section, preferably in longitudinal section with respect to the longitudinal center axis. The imaginary level should continuously accommodate the wing transition point. In addition, it is particularly preferably perpendicular to the longitudinal center axis. The straight line is in turn perpendicular to the imaginary plane and in this respect is preferably arranged parallel to the longitudinal center axis of the wing.

Immediately at the wing transition point, the second profile surface encloses an angle of at least 0° and at most 60° with the straight line. Provision can therefore be made for the second profile surface to merge into the second profile surface parallel to the straight line. However, the angle is preferably greater than 0°. For example, it is at least 15° and at most 60°, at least 30° and at most 50° or approximately or exactly 0°, approximately or exactly 30° or approximately or exactly 50°. In the stated angle range, a premature detachment of the air film from the second profile surface is reliably prevented, so that the resulting free jet air flow forms the support vortex particularly effectively.

A development of the invention provides that the deflection element can be displaced for the global and/or local change in a flow cross-sectional area of the air outlet gap. In this respect, the deflection element can be displaced in such a way, in particular by means of a control drive, that the size of the air outlet gap is changed, namely either globally and/or locally. The global change in the air outlet gap or the flow cross-sectional area of the air outlet gap means that the size of the air outlet gap or the flow cross-sectional area changes uniformly over the entire extent of the air outlet gap, i.e. increases or decreases. The local change, on the other hand, means an increase or decrease in the air outlet gap or the cross-sectional area of flow only in certain areas. For example, for local change, the deflection element is displaced in such a way that the air outlet gap is increased in some areas and reduced in some areas. By changing the flow cross-sectional area of the air outlet gap, the missile, in particular a flight direction of the missile, can be controlled in a simple manner.

A further development of the invention provides that the distance between the deflection element and the second profile surface can be changed uniformly in order to globally change the flow cross-sectional area of the air outlet gap. Uniform changing means a uniform enlargement or a uniform reduction of the air outlet gap. For example, the deflection element is shifted parallel to the longitudinal central axis of the wing, namely away from the second profile surface to increase the air outlet area and towards the second profile surface to reduce the flow cross-sectional area. This enables the missile to be controlled particularly effectively by adjusting the vortex intensity of the support vortex.

A development of the invention provides that the deflection element can be tilted with respect to the second profile surface in order to locally change the flow cross-sectional area of the air outlet gap. By tilting the deflection element, the air outlet gap is changed locally, in particular partially enlarged and partially reduced. The tilting takes place, for example, with respect to Central longitudinal axis of the wing. The deflection element is preferably designed in such a way that when the deflection element is aligned parallel to the longitudinal center axis and to this extent at an angle of 0°, the air outlet gap has a consistently constant size in the circumferential direction with respect to the longitudinal center axis. If the angle changes, on the other hand, there is a local change in the flow cross-sectional area. In turn, such a configuration enables a particularly efficient control of the missile.

A development of the invention provides that first control elements emanate from the first profile surface and/or second control elements emanate from the second profile surface, each of which has a control fin. The first control elements and the second control elements are used to control the missile, namely by influencing the air film and/or the suspension vortex. It can be provided that only the first control elements but not the second control elements are present, or only the second control elements but not the first control elements. However, it can also be provided that both the first control elements and the second control elements are implemented. Each of the control elements has a control fin, which is designed, for example, like a plate or like an aerofoil. In the latter case, the control fin can be chordwise symmetrical or have an aerodynamic profile. The first control elements, which emanate from the first profile surface, are used to directly influence the support vortex, whereas the second control elements, which emanate from the second profile surface, directly influence the air film and thus only indirectly influence the support vortex. Depending on the angle of attack of the control elements, the support vortex and/or the air film are deflected, so that the missile can be controlled in an effective manner. The control elements can be implemented in addition to or as an alternative to the displaceability of the deflection element. The second control elements are arranged, for example, in the second area of the second profile surface or are mounted in this.

A development of the invention provides that the first control elements and/or the second control elements are coupled in terms of drive technology to a control drive of the missile via a common coupling element. The control drive is used to adjust the control elements. In terms of drive technology, it is only indirectly connected to the control elements via the common coupling element. For this purpose, the control elements on the one hand and the control drive on the other act on the coupling element. In particular, the control drive acts on the coupling element at a distance from the control elements. This allows the control elements to be adjusted simultaneously by means of the control drive. The deflection element can also be additionally or alternatively connected to the coupling element.

A development of the invention provides that the coupling element is coupled to the first control elements and/or the second control elements and/or the control drive via a ball joint and a lever arm. To this extent, each control element and/or the control drive is assigned a ball element and a lever arm, by means of which they are connected to the control drive in terms of drive technology. The use of the ball joint ensures extremely flexible adjustment of the controls using the control drive.

A development of the invention provides that the coupling element is designed as a control ring. The control ring preferably encompasses the longitudinal central axis of the wing continuously and completely in the circumferential direction. It acts on both the first control elements and the second control elements - if present - in order to couple them to the control drive. The control ring is arranged in such a way that it can not only describe a rotational movement in the circumferential direction with respect to the longitudinal center axis, but that it can also be tilted and is therefore present in the manner of a swash plate. This enables the already mentioned flexible actuation of the control elements by means of the control drive.

A further development of the invention provides that the control drive has a plurality of actuators which are each spaced apart from one another in terms of drive technology and are coupled to the coupling element. Preferably, the actuators are evenly spaced from each other, so that in the case of two actuators, they act on the coupling element at a distance of 180°, in the case of three actuators at a distance of 120° and in the case of four actuators at a distance of 90° . The use of multiple actuators allows the coupling element to be displaced not only in the circumferential direction with respect to the longitudinal center axis, but also in the radial direction, so that the flexible adjustment of the control elements already described is implemented.

The invention also relates to a method for operating a missile, in particular a missile according to the statements made within the scope of this description. It is provided that the missile has a wing, seen in section on the one hand from a lower first profile surface during normal operation of the missile and on the other hand from an upper ren, is delimited at a wing transition point with the second profile surface converging with the first profile surface, the first profile surface enclosing at least one air inlet opening and the second profile surface enclosing at least one air outlet opening and the missile has a propulsion device with an air conveying device which is used to suck in air through the at least one air inlet opening and is operated for discharging the sucked-in air through the at least one air outlet opening, wherein the at least one air outlet opening is overlapped at least in regions by a deflection element which, with the second profile surface, delimits an air outlet gap that is flow-connected to the air outlet opening, so that the air is parallel to the second profile surface is deployed.

Reference has already been made to the advantages of such a configuration of the missile or such a procedure. Both the missile and the method for operating it can be further developed in accordance with the statements made within the scope of this description, so that reference is made to them in this respect.

A further development of the invention provides that, by means of the drive device, air is drawn in in the form of an intake air flow in an intake direction through the air inlet opening and is discharged through the air outlet gap in an exit direction which is angled with respect to the intake direction in such a way that the air forms a free jet air flow in a free jet direction at the wing transition point , so that a support vortex is formed between the intake air flow and the free jet air flow, which vortex is present at least in regions below the first profile surface. This too has already been addressed.

• 1 eine schematische Darstellung einer ersten Ausführungsform eines Flugkörpers im Längsschnitt bezüglich einer Längsmittelachse, sowie
• 2 eine vereinfachte schematische Darstellung des Flugkörpers in einer zweiten Ausführungsform.

The invention is explained in more detail below with reference to the exemplary embodiments illustrated in the drawing, without the invention being restricted. while showing

• 1 a schematic representation of a first embodiment of a missile in longitudinal section with respect to a longitudinal central axis, and
• 2 a simplified schematic representation of the missile in a second embodiment.

the 1 shows a schematic representation of a missile 1 in a first embodiment, which has at least one wing 2, which is round or ring-shaped in cross section with respect to a longitudinal center axis 3 of the wing. The wing 2 is delimited in the axial direction with respect to the longitudinal central axis 3 in a first direction by a first profile surface 4 and in a second direction by a second profile surface 5 . Each of the profile surfaces 4 and 5 is itself ring-shaped. The two profile surfaces 4 and 5 converge at a wing transition point 6, the transition between the profile surfaces 4 and 5 being preferably discontinuous at this point as seen in section. The profile surfaces 3 and 5 thus preferably together form an edge at the wing transition point 6 . For this purpose, the wing transition point 6 is present, for example, in the form of a geometric circle into which the first profile surface 4 opens at a first angle and the second profile surface 5 at a second angle different from the first angle.

The first profile surface 4 includes an air inlet opening 7 and the second profile surface 5 an air outlet opening 8 . The air inlet opening 7 and the air outlet opening 8 are fluidically connected to one another via a flow channel 9 , a ducted propeller 10 of an air conveying device 11 being rotatably mounted in the flow channel 9 . The air conveying device 11 is part of a drive device 12 of the missile 1. It can be clearly seen that the first profile surface 4 is on a side of the missile 1 facing a floor 13, whereas the second profile surface 5 faces away from the floor 13. The missile 1 is shown here during normal and intended flight operations.

The air inlet opening 7 and the air outlet opening 8 are located in the center with respect to the longitudinal central axis 3 and are bordered by the first profile surface 4 or the second profile surface 5 on an inner side 14 of the wing in the radial direction. The flow channel 9, which connects the air inlet opening 7 and the air outlet opening 8 to one another in terms of flow, is present, for example, in the form of a cylinder, in particular a circular cylinder and particularly preferably in the form of a right circular cylinder. The air inlet opening 7 and the air outlet opening 8 particularly preferably have the same flow cross section. The air inlet opening 7 is arranged on an underside and the air outlet opening 8 on an upper side of the missile 1 . The air outlet opening 8 is overlapped at least in regions, in the exemplary embodiment shown here completely, by a deflection element 15 which, together with the second profile surface 5 , delimits an air outlet gap 16 . Viewed in longitudinal section, the air outlet gap 16 is therefore located between the deflection element 15 and the second profile surface 5 .

The air outlet gap 16 is fluidically connected to the air outlet opening 8 via a connecting channel 17 , the connecting channel 17 also being delimited at least in regions by the deflection element 15 and the second profile surface 5 . In the exemplary embodiment shown here, both the air outlet gap 16 and the connecting channel 17 are designed to be continuous in the circumferential direction with respect to the longitudinal center axis 3 and each completely enclose the longitudinal center axis 3 . According to the exemplary embodiment shown, a flow cross section of the connecting channel 17 decreases in the direction of the air outlet gap 16, so that the connecting channel 17 is present in the manner of a nozzle and the air outlet gap 16 represents an orifice of this nozzle.

The deflection element 15 can be displaced in such a way that a flow cross-sectional area of the air outlet gap 16 can be changed locally and/or globally. For this purpose, the deflection element 15 is coupled in terms of drive technology, for example, to a control drive of the missile 1 (not shown here). Additionally or alternatively, several first control elements 18 are connected to the control drive in terms of drive technology. In addition or as an alternative to the first control elements 18, second control elements (not shown here) can be present. The first controls 18 are based on the first profile surface 4; the second control elements from the second profile surface 5. The control elements 18 each have a control fin 19 and are each drive-connected via a lever arm, not shown here, and a ball joint, also not shown, to a coupling element 20, which is present, for example, as a control ring. The coupling element 20 is used to drive the connection of the control elements 18 to the control drive, in particular to several actuators of the control drive, which are not shown here.

When the missile 1 is in flight, the propulsion device 12 is operated in such a way that air is conveyed from the underside of the missile 1 in the direction of its upper side or in such a way that air in the form of an intake air flow 21 is sucked in in an intake direction through the air inlet opening 7 and through the flow channel 9 is conveyed in the direction of the air outlet opening 8. The conveyed air is discharged through the air outlet opening 8 and supplied to the air outlet gap 16 through the connecting duct 17 . The air is discharged through the air outlet gap 16 in an outlet direction in the form of an air film 22 in such a way that the air film 22 bears against the second profile surface 5 . For this purpose, the air film 22 is preferably discharged at the air outlet gap 16 parallel to the second profile surface 5 .

Due to the Coandä effect, the air film 22 runs along the second profile surface 5 up to the wing transition point 6. There the air film 22 detaches from the wing 2 and flows again as a free jet air flow 23 onto the underside of the missile 1, namely in a free jet direction. Here, the free jet air flow 23 excites a support vortex 24 which provides at least part of the lift for the missile 1 . The support vortex 24 is intensified by the intake air flow 21. It can be seen that both the free jet air flow 23 and the intake air flow 21 are each tangential to the support vortex 24. Due to the annular shape of the wing 2, the support swivel 24 formed has a toroidal shape. For the sake of clarity, the support swivel 24 is only shown on one side of the wing 2 . Of course, the support swivel 24 preferably completely and continuously surrounds the longitudinal center axis 3 in the circumferential direction. Due to the high speed of the air film 22, this causes a suction effect on the surrounding air. This means that more air is supplied to the air film 22 when it flows over the second profile surface 5 . This is indicated by the arrows 25.

In order to achieve precise positioning of the support vortex 24 below the wing 2 , a turbulence chamber 26 is formed on the wing 2 , which is delimited outwards in the radial direction by the wing transition point 6 . The vortex chamber 26 is formed, for example, by a curvature or a recess in the first profile surface 4 . In order to achieve a particularly advantageous flow around the second profile surface 5 by the air film 22, the second profile surface 5 merges at the wing transition point 6 at an angle α into the first profile surface 4, which is determined with respect to a straight line 27 which runs parallel to the longitudinal center axis 3 runs through the wing transition point 6. The straight line 27 is in particular perpendicular to an imaginary straight line which takes up the wing transition point 6 continuously.

It should be pointed out that the configuration of the support surface 2 as a ring-shaped support surface 2 described here merely represents an advantageous exemplary embodiment. The wing 2 can also be basically straight and in this case, for example, have the illustrated cross section. In principle, the advantages described can also be achieved with such an embodiment.

the 2 shows a second embodiment of the missile 1 in a schematic and simplified manner, with some elements not being shown for reasons of clarity. In principle, the second embodiment corresponds to the first one management form, so that the corresponding explanations are referred to and only the differences are discussed below. On the one hand, these are due to the fact that the angle α is negative or greater than 180°. For example, the angle α is at least 270° and less than 360°. Alternatively, it is at least 300° and at most 345° or at most 330°. This means that the second profile surface 5 approaches the longitudinal center axis 3 again, starting from the outside of the wing where it is at the greatest distance from the longitudinal center axis 3, before it transitions into the first profile surface 4 at the wing transition point 6.

A further difference lies in the fact that the air inlet opening 7 is clearly spaced in the axial direction with respect to the longitudinal central axis 3 from an imaginary plane which continuously accommodates the wing transition point 6 . For example, the distance is at least 20%, at least 30%, at least 40% or at least 50% of an axial distance between the plane and a point of the second profile surface 5 or the air outlet opening that is furthest away from the plane when viewed in section in the axial direction. As a result, the air inlet opening 7 is arranged at the appropriate distance away from the plane, so that the formation of the support vortex is facilitated.

Another difference can be seen in the fact that the second profile surface 5 is not continuously curved, but is composed of a first area 28 and a second area 29 . Seen in section, the first region 28 extends, starting from the wing transition point 6, directly to the second region 29. It is preferably continuously curved, seen in section. The second area 29 , on the other hand, is flat when viewed in section or has a different curvature than the first area 28 . For example, the transition between the first area 28 and the second area 29 is discontinuous when viewed in section. Alternatively, however, it can also be continuous.

It should be noted that each of the noted differences is individually applicable to the first embodiment. It is therefore not necessary for the differences to always occur in combination with one another.

In both embodiments, the missile 1 described has the advantage that it works in an extremely energy-efficient manner due to the use of the support vortex 24 to provide at least part of the lift. In addition, extremely precise control of the missile 1 by means of the displaceable deflection element 15 and/or the control elements 18 is possible. In particular, the missile 1 can hover in the air analogously to a helicopter. Nevertheless, it can achieve quite high speeds because, unlike the helicopter, it is not limited by a maximum flow speed at the blade tips of a rotor.

Reference List

1

missile

2

wing

3

longitudinal central axis

4

profile surface

5

profile surface

6

wing transition point

7

air intake opening

8th

air outlet opening

9

flow channel

10

ducted propeller

11

air conveyor

12

drive device

13

floor

14

wing inside

15

deflection element

16

air outlet gap

17

connecting channel

18

first controls

19

control fin

20

coupling element

21

intake airflow

22

air film

23

free jet air flow

24

lifting swivel

25

Arrow

26

vortex chamber

27

Just

28

first area

29

second area

Claims (11)

Missile (1), with a wing (2), seen in section, on the one hand by a lower first profile surface (4) when the missile (1) is operated as intended and on the other hand by an upper one at a wing transition point (6) is delimited with the second profile surface (5) converging with the first profile surface (4), the first profile surface (4) enclosing at least one air inlet opening (7) and the second profile surface (5) enclosing at least one air outlet opening (8) and the missile ( 1) has a drive device (12) with an air conveying device (1), which is provided and designed to suck in air through the at least one air inlet opening (7) and to discharge the sucked-in air through the at least one air outlet opening (8), the at least an air outlet opening (8) is overlapped at least in regions by a deflection element (15) which, with the second profile surface (5), delimits an air outlet gap (16) which is in flow connection with the air outlet opening (8), and the air outlet opening (8) with the air outlet gap ( 16) via a connecting channel (17) is fluidically connected, seen in section between the second profile surface (5) and the order steering element (15), the second profile surface (5) having a first area starting from the wing transition point (6) and a second area adjoining the first area and delimiting the air outlet gap (16), wherein - seen in section - the first area is curved and the second area is curved or flat, and wherein the connecting duct (17) has a cross section that decreases in the direction of the air outlet gap (16), so that it is designed in the manner of a nozzle, with a shape of the connecting duct (17) such is selected so that the air in the air outlet gap (16) has a flow speed in the subsonic range, characterized in that the at least one air inlet opening (7) and the at least one air outlet opening (8) are formed via a section centrally in the wing (2). Flow channel (9) are fluidically connected, wherein in the flow channel (9) a jacket propeller (10) of the air conveying device (11) is arranged to be rotatable about an axis of rotation. missile after claim 1 , characterized in that the at least one air inlet opening (7) is arranged centrally in the first profile surface (4) and/or the at least one air outlet opening (8) is arranged centrally in the second profile surface (5). Missile according to one of the preceding claims, characterized in that the wing (2) is ring-shaped with respect to a longitudinal central axis (3) and the profile surfaces (4, 5) are axially spaced apart from one another in certain areas when viewed in longitudinal section. Missile according to one of the preceding claims, characterized in that the first profile surface (4), viewed in section, is set back in some areas in relation to the wing transition point (6) in the direction of the second profile surface (5), so that a vortex chamber (26) is formed, which is Wing transition point (6) is embraced. Missile according to one of the preceding claims, characterized in that the first profile surface (4) and the second profile surface (5) are curved in the same direction, at least in certain areas. Missile according to one of the preceding claims, characterized in that the second profile surface (5) at the wing transition point (6) merges into the first profile surface (4) at an angle which is perpendicular to an imaginary plane continuously accommodating the wing transition point (6). standing straight line (27) is at least 0° and at most 60°. Missile according to one of the preceding claims, characterized in that the deflection element (15) can be displaced for the global and/or local change in a flow cross-sectional area of the air outlet gap (16). Missile according to one of the preceding claims, characterized in that first control elements (18) emanate from the first profile surface (4) and/or second control elements emanate from the second profile surface (5), each having a control fin (19). Missile according to one of the preceding claims, characterized in that the first control elements (18) and/or the second control elements are drivingly coupled to a control drive of the missile (1) via a common coupling element (20). Method for operating a missile (1) according to one or more of the preceding claims, wherein the missile (1) has a wing (2) which, viewed in section, extends on the one hand from a lower first profile surface (4) when the missile (1) is operated as intended. and on the other hand is delimited by an upper second profile surface (5) which converges with the first profile surface (4) at a wing transition point (6), the first profile surface (4) having at least one air inlet opening (7) and the second profile surface (5) having at least one air outlet opening (8) and the missile (1) has a drive device (12) with an air conveying device (11) which is used for sucking in air through the at least one air inlet opening (7) and for discharging the sucked air through the at least one air outlet opening (8 ) is operated, the at least one air outlet opening (8) being overlapped at least in regions by a deflection element (15) which, with the second profile surface (5), delimits an air outlet gap (16) which is in flow connection with the air outlet opening (8), so that the air flows parallel to the second profile surface (5) is discharged, and the air outlet opening (8) is fluidically connected to the air outlet gap (16) via a connecting duct (17) which, viewed in section, is present between the second profile surface (5) and the deflection element (15), the second profile surface (5) has a first area starting from the wing transition point (6) and a second area adjoining the first area and delimiting the air outlet gap (16), wherein - viewed in section - the first area is curved and the second area is curved or is planar, and wherein the connecting duct (17) decreases in the direction of the air outlet gap (16). n cross-section, so that it is configured like a nozzle, the shape of the connecting duct (17) being chosen such that the air in the air outlet gap (16) has a flow rate in the subsonic range, characterized in that the at least one air inlet opening ( 7) and the at least one air outlet opening (8) are fluidically connected via a flow channel (9) formed centrally in the wing (2) in section, with a ducted propeller (10) of the air conveying device (11) revolving around an axis of rotation in the flow channel (9). is rotatably arranged. procedure after claim 10 , characterized in that by means of the drive device (12) air in the form of an intake air flow (21) is sucked in in an intake direction through the air inlet opening (7) and is discharged through the air outlet gap (16) in an outlet direction which is angled with respect to the intake direction in such a way that at the Wing transition point (6) where the air forms a free-jet air flow (23) in a free-jet direction, so that a support vortex (24) is formed between the intake air flow (21) and the free-jet air flow (23), which is at least partially below the first profile surface (4).

Images