FUSELAGE FOR AN AIRCRAFT WITH FUSELAGE-INTEGRATED TAILPLANE

Abstract

A fuselage for an aircraft. The fuselage has a control element with an integrated engine outlet. The control element is integrated at a rear end of the fuselage, so that the control element terminates flush with an outer skin of the fuselage in a circumferential direction of the fuselage. An outer wall of the control element surrounds the engine outlet wherein the engine outlet is directed towards an open rear side of the control element. The control element is connected to the fuselage such that the control element jointly the engine outlet is pivotable about a rotation axis with respect to the fuselage. The rotation axis runs transversely to a longitudinal direction of the fuselage and the control element functions as a tailplane when pivoting about the rotation axis.

Claims

1. A fuselage for an aircraft, the fuselage comprising: a control element with an integrated engine outlet; wherein the control element is integrated at a rear end of the fuselage and is a rearward extension of the fuselage in a longitudinal direction of the fuselage, so that the control element terminates flush with an outer skin of the fuselage in a circumferential direction of the fuselage; wherein an outer wall of the control element surrounds the engine outlet, such that the engine outlet is directed towards an open rear side of the control element; wherein the fuselage transitions into the control element at a transition point, the fuselage and the control element having substantially similar cross-sections at the transition point and, at the transition point, a width of the control element is substantially similar to a width of the fuselage and a height of the control element is substantially similar to a height of the fuselage; wherein the control element is connected to the fuselage such that the control element is pivotable together with the engine outlet about a rotation axis with respect to the fuselage; and wherein the rotation axis runs transversely to the longitudinal direction of the fuselage and the control element functions as a tailplane when pivoting about the rotation axis.

2. The fuselage according to claim 1, wherein the engine outlet is a thrust vectoring nozzle.

3. The fuselage according to claim 1, wherein the engine outlet has a convergent nozzle portion and a divergent nozzle portion arranged therebehind in the longitudinal direction.

4. The fuselage according to claim 3, wherein the convergent nozzle portion has two control blades which are adjustable relative to each other so that a cross-section of the convergent nozzle portion is variable.

5. The fuselage according to claim 3, wherein: the divergent nozzle portion has two control blades; the two control blades of the divergent nozzle portion are adjustable relative to each other, so that a cross-section of the divergent nozzle portion is variable; or the two control blades of the divergent nozzle portion are adjustable jointly, so that a cross-section of the divergent nozzle portion remains same and an outlet direction of the engine outlet is changed.

6. The fuselage according to claim 5, wherein the cross-section of the convergent nozzle portion is variable independently of the cross-section of the divergent nozzle portion.

7. The fuselage according to claim 5, wherein the two control blades of the divergent nozzle portion are angled and each of the two control blades of the divergent nozzle portion meets the outer wall of the control element at an angle different from 90°.

8. The fuselage according to claim 5, wherein a rear edge of the two control blades of the divergent nozzle portion is serrated.

9. The fuselage according to claim 5, wherein: the control element has a lateral nozzle cover on both sides; and each lateral nozzle cover is connected to a respective one of the two control blades of the divergent nozzle portion and is entrained by the respective one of the two control blades of the divergent nozzle portion when the respective one of the two control blades of the divergent nozzle portion moves.

10. The fuselage according to claim 1, wherein the engine outlet is double-walled, at least in portions, so that a cooling flow can be guided through the double-walled portion.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0042] Further details are described with reference to the figures. The figures are schematic and not to scale.

[0043] FIG. 1 shows a schematic representation of an aircraft.

[0044] FIG. 2 shows a schematic representation of an aircraft with a fuselage constructed according to the principles described herein.

[0045] FIG. 3 shows a schematic representation of a control element with integrated engine outlet.

[0046] FIG. 4 shows a schematic representation of a control element with integrated engine outlet, wherein the control element is shown in various positions.

[0047] FIG. 5 shows a schematic representation of the convergent and divergent nozzle portion of the engine outlet.

[0048] FIG. 6 shows a schematic representation of the convergent portion of the engine outlet.

[0049] FIG. 7 shows a schematic representation of the engine outlet with a cavity for cooling air flowing past.

[0050] FIG. 8 shows a schematic representation of the engine outlet with a thrust reversal device.

DETAILED DESCRIPTION

[0051] FIG. 1 shows an aircraft 1 in the form of a jet aircraft with two engines. It should be understood that the reference to an aircraft always means that the corresponding statement can also refer to an aircraft in general.

[0052] The aircraft 1 has a fuselage 3 with wings 5 arranged laterally thereon. In addition, the aircraft 1 also has control surfaces (tailplane, vertical stabilizer, landing flaps, etc.) which are arranged on the fuselage or the wings. The vertical stabilizers 12 and the tailplane 13 on the rear edge of the wings 5 are shown here by way of example.

[0053] Air inlet openings 7 are arranged on the fuselage, typically laterally on the fuselage and below the wings 5. However, it should be understood that the positioning of the air inlet openings is only shown here as an example and is not decisive for the design of the described air inlet device.

[0054] The air inlet openings 7 draw in air from the environment and pass it on to, among other things, the engine 10 or engines 10. The air is guided from the air inlet opening 7 via a duct to the engine 10 or its first compressor stage.

[0055] In order to keep the radar signature of an aircraft low, various measures are sometimes taken. One of these is to avoid a direct line of sight, from the front, of the engine and its first compressor stage, because the engine or its first compressor stage is a very strong reflector for radar signals. Such measures concerning the air inlet opening 7 serve as camouflage against radar reconnaissance from the front. Other measures include shaping the outer contour of the aircraft according to certain design principles (long edges, parallel edges and surfaces, as already described above).

[0056] The focus of this description is on the design of the tail of the aircraft 1, which is such as to reduce a radar signature of the tail and engine outlets.

[0057] FIG. 1 shows an aircraft with two engines 10. It can be seen that the upper surface of the fuselage is a multi-curved surface, resulting from the presence of the two engines and two engine outlets at the tail of the aircraft.

[0058] FIG. 2 shows an aircraft 1 with a fuselage 3 of which the rear end 2 is designed according to the principles described here.

[0059] A control element 20 with an integrated thrust vectoring nozzle adjoins the rear end of the fuselage 3. The aircraft 1 is basically similar in construction to the aircraft 1 from FIG. 1. FIG. 2 differs from FIG. 1 only in the use of the control element 20 and some further adjustments to the shape of the fuselage 3. While the surface of the fuselage 3 in FIG. 1 assumes a multiple curved shape at the top side of the fuselage to transition into the shape of the engine outlets 10, in FIG. 2 the fuselage 3 has a nearly flat and at most slightly curved upper surface at the outer skin 4. Thus, by using the control element 20, the outer skin 4 can be designed to be largely flat and level instead of containing multiple curved portions as in FIG. 1. Similarly, the fuselage 3 of FIG. 2 can be made flat and level on the underside of the fuselage, as also shown and described on the outer skin 4 for the upper surface.

[0060] The vertical stabilizer 12 guides an airflow along the upper surface of the fuselage to the control element 20, so that the control element with its upper control surface can function as a tailplane and generate a moment about a transverse axis of the fuselage (so-called pitching motion). The same applies to the lower surface of the fuselage, which guides an airflow to the lower control surface of control element 20. Thus, the control element 20 has a good aerodynamic effect when used as a tailplane and deflected from its initial state shown in FIG. 2.

[0061] The wings 5 can have further control surfaces 13 at their trailing edge. These control surfaces 13 can be used as landing flaps, for example. However, the control surfaces 13 can also additionally function as a tailplane if a stronger moment about the transverse axis of the fuselage 3 must be generated in a special manoeuvre. Furthermore, the control surfaces 13 can be used to generate a moment about the longitudinal axis of the fuselage (so-called rolling motion), for example by deflecting a control surface 13 on one wing upwards and a control surface of the other wing downwards. FIG. 2 also shows that the outer contour of the fuselage transitions seamlessly into the outer contour of the control element 20 and that the control element 20 has very few edges and surfaces visible from the outside (compared to a round nozzle with a plurality of plates arranged movably in the circumferential direction of the round nozzle).

[0062] FIG. 3 shows a rear view of the control element 20. The control element 20 has an outer wall 40 which surrounds the elements of the control element in the circumferential direction. The control element 20 has an upper control surface 21 and a lower control surface 22. The engine outlets 23 can be seen as rear openings of the engine nozzle. The engine outlets 23 are each laterally bounded by two nozzle control blades 24. The nozzle control blades 24 are pivotable so that a cross-section of the rear opening of the engine outlets 23 can be changed by moving the nozzle control blades 24 towards or away from each other.

[0063] The nozzle control blades 24 are angled and their upper and lower edges meet the upper and lower portions of the outer wall 40 at an angle different from 90°. In particular, the nozzle control blades 24 are angled at such an angle that they correspond to the angle of the lateral portion of the outer surface of the outer wall to comply with the design principle of the edges being parallel to each other.

[0064] FIG. 4 shows a side view of the control element 20 and a portion of the engine 10 that transitions into the control element 20. The fuselage surrounding the engine 10 is not shown in this representation. FIG. 4 shows the control element 20 in three different states A, B, C.

[0065] The initial state of the control element 20 can be referred to as state B. The control element 20 is oriented so that it is parallel to the longitudinal direction 8 or longitudinal axis of the fuselage. In this state, no moment about the transverse axis is generated by the control element 20 during operation of the aircraft.

[0066] From the initial state B, the control element 20 can be moved upwards (state A) or downwards (state C), more specifically with the aid of an actuator 25, which receives an actuation signal from a flight control computer, wherein the actuation signal is output from an autopilot or a control element operated by a human pilot. The actuator 25 applies a force to the control element 20 so that the control element 20 performs a pivoting movement about the rotation axis 26.

[0067] Although only one actuator 25 is shown in FIG. 4, it should be understood that, for redundancy reasons, more than one actuator together with associated mechanism may be provided in order to change the position of the control element 20.

[0068] In state A, the air flowing over the upper control surface 21 generates a moment which rotates the fuselage about its transverse axis. When the control element 20 is moved to state A, the upper surface 21 slides on the fuselage side under a shell surface 27 and the lower control surface is still covered by the shell surface 27, so that no opening is created in the outer skin of the aircraft when the control surface 20 is moved from the initial state B to a deflected state A, C. The same applies for the upper and lower control surfaces of control element 20 in state C.

[0069] The shell surface 27 can be part of the fuselage 3 or part of the control element 20. In any case, the shell surface 27 does not move about the rotation axis 26, but is static with respect to the fuselage. The upper and lower control surfaces 21, 22 of the control element slide under the shell surface or are pulled out from under it when the control element 20 performs its pivoting movement about the rotation axis 26.

[0070] In FIG. 4, the lateral portions of the outer wall are not visible so as to allow a view of the actuator 25, the rotation axis 26 and the nozzle control blades 24. However, it should be understood that these elements are hidden by the outer wall when the aircraft is in operation.

[0071] FIG. 5 shows a schematic representation of the engine outlet in plan view with the convergent nozzle portion 28 and the divergent nozzle portion 29. The engine outlet is covered laterally by the nozzle covers 30.

[0072] In FIG. 5, two engine outlets are shown, similarly to FIGS. 2 and 3. It should be recognized, however, that the principles described here can be used for a single engine or multiple engines.

[0073] A control element 20 as described herein can also be provided separately for one engine in each case. If these two engines are at a certain distance from a central axis of the fuselage, then by two control elements, which are separated and spaced from each other, a torque about the longitudinal axis (roll) of the fuselage can also be generated via a separate actuation, for example by deflecting one control element upwards and the other control element downwards.

[0074] With reference to the upper engine in FIG. 5, the details of the convergent nozzle portion 28 and of the divergent nozzle portion 29 are explained. The convergent nozzle portion 28 is formed by two mutually opposed control blades 31. These taper conically towards the outlet opening (to the right), i.e. the cross-section of the nozzle is reduced in the convergent nozzle portion. An actuator 34 is connected to the control blades 31 via connecting rods and is designed to move the control blades 31 so that a ratio of the cross-section A.sub.7 (inlet side of the convergent nozzle portion) to the cross-section A.sub.8 (outlet side of the convergent nozzle portion) can be varied. In addition, an actuator 35 is connected to the control blades 32 of the divergent nozzle portion 29 and is configured to move the control blades 32 in order to vary a ratio of the cross-section A.sub.8 (inlet side of the divergent nozzle portion) to the cross-section A.sub.9 (outlet side of the divergent nozzle portion).

[0075] The actuator 35 for the divergent nozzle portion 29 is connected to the control blades 31 and the outer wall of the control element via connecting rods 33. Thus, the control blades 32 of the divergent nozzle portion 29 are entrained when the control blades 31 of the convergent nozzle portion 28 are moved.

[0076] Even though the control blades 32 of the divergent nozzle portion 29 in FIG. 5 run parallel to each other, these control blades 32 can assume different positions with respect to each other and with respect to the control element. The control blades 32 can be brought into a position such that the cross-section A.sub.9 is larger than the cross-section A.sub.8. The control blades 32 of both engine outlets can both be pivoted laterally at the same time in the same direction or in different directions (in the illustration of FIG. 5 upwards or downwards, i.e. in the direction of a nozzle cover). If both control blades 32 of both engine outlets are pivoted jointly and simultaneously in the direction of the upper or lower nozzle cover 30, this generates a moment about the vertical axis of the aircraft (yaw movement). If the two control blades 32 of the upper engine outlet are swung out laterally in a first direction (away from the central axis of the fuselage) and the two control blades 32 of the lower engine outlet are swung out laterally in the other direction, opposite to the first direction, such that a distance 38 between the inner control blades 32 is increased or maximized, a braking effect can be generated. The control surfaces 32 of the divergent nozzle portion 29 can also be deflected so that a distance 38 between the inner control blades 32 becomes minimal. This reduced distance between the two engine outlets also reduces the radar signature.

[0077] FIG. 6 shows a schematic isometric representation of the control blades 31 of the convergent nozzle portion 29. In the example of FIG. 6, two pairs of control blades 31 are shown, wherein one pair of control blades 31 is associated with one power unit in each case. The control blades 31 of an engine are moved towards each other by the actuator 34 via a linkage or a gearing in such a way that the ratio of the cross-sections A.sub.7 and A.sub.8 shown in FIG. 5 is varied. The actuators 34 are controlled by a flight control computer.

[0078] All actuators 25, 34, 35 described herein can be designed as hydraulic, electric or electrohydraulic drives. The actuators generate a movement that is transmitted to an element to be moved by a mechanism, for example in the form of a linkage and/or gearing. The actuators can perform a linear movement or a rotary movement.

[0079] FIG. 7 shows a schematic cross-sectional representation of the engine 10 and of the control element 20. In FIG. 7 the control element 20 is shown without deflection about the rotation axis, i.e. in the same state as in FIG. 4, B. This means that the upper control surface 21 and the lower control surface 22 do not deflect air flowing along the outer skin of the fuselage and thus do not generate a moment about the transverse axis of the aircraft.

[0080] As can be seen from FIG. 7, a trailing edge of the control blades 32 of the divergent nozzle portion is zig-zag-shaped and does not form a single, continuous, straight breakaway edge. This increases the effective length of the entire trailing edge and also ensures that the hot exhaust gases from the engine are better mixed with the ambient air to reduce a heat signature of the aircraft.

[0081] The outer wall of the engine 10 is double-walled with a cooling air duct 36. In the cooling air duct 36, cool air from the environment flows along the longitudinal axis of the engine towards the engine outlet. This air cools the engine. In the region of the convergent nozzle portion and the associated control blades, the cooling air from the cooling air duct 36 flows inwards, for example through openings in the inner wall of the cooling air duct. The cool air from the cooling air duct can mix with the hot exhaust gases of the engine at this point and can cool the exhaust gases. Due to the pressure and flow conditions in the engine outlet, the air exiting the cooling air duct flows close to the wall of the engine outlet towards the outlet opening. Therefore, the cool air also flows along the control blades 32 of the divergent nozzle portion and cools them as well.

[0082] FIG. 8 shows another function of the engine 10. In state A, the engine is shown as also shown in FIG. 7 above, i.e. without deflection of the control element 20 about the rotation axis.

[0083] However, the engine 10 is extended by two inner deflector plates 37A (and an associated drive, for example in the form of an actuator/motor together with gearing and linkage) which can be moved towards each other from below and above to redirect the engine exhaust flow so as to effect a thrust reversal.

[0084] The thrust reversal state is shown in FIG. 8 at the bottom, marked with the letter B. The inner deflector plates 37A are pivoted one downwards and one upwards and rest against each other in such a way that they block the path of the engine exhaust gases to the rear outlet opening and redirect the exhaust gases upwards and downwards respectively, as illustrated by the two arrows.

[0085] When the inner deflector plates 37A are pivoted from state A of FIG. 8 to state B of FIG. 8, openings on the outer walls of the engine are released by pivoting outer deflector plates 37B from a closed position (state A) to an open position (state B), releasing the exhaust gases from the aircraft against the direction of flight. The openings closed or released by the outer deflector plates 37B are located in front of the inner deflector plates 37A. Thus, the deflector plates 37A, 37B effect a thrust reversal.

[0086] While at least one example embodiment of the invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the example embodiment(s). In addition, in this disclosure, the terms “comprise” or “comprising” do not exclude other elements or steps, the terms “a”, “an” or “one” do not exclude a plural number, and the term “or” means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. Reference signs in the claims are not to be regarded as a limitation.

LIST OF REFERENCE SIGNS

[0087] 1 Aircraft [0088] 2. Rear end [0089] 3 Fuselage [0090] 4 Outer skin [0091] 5 Wings [0092] 6 Rear side [0093] 7 Air inlet opening [0094] 8 Longitudinal direction [0095] 9 Air inlet duct [0096] 10 Engine [0097] 12 Vertical stabilizer [0098] 13 Control surfaces [0099] 20 Control element with integrated thrust vectoring nozzle [0100] 21 Upper control surface [0101] 22 Lower control surface [0102] 23 Engine outlet, engine nozzle [0103] 24 Nozzle control blade [0104] 25 Actuator [0105] 26 Rotation axis [0106] 27 Shell surface [0107] 28 Convergent nozzle portion [0108] 29 Divergent nozzle portion [0109] 30 Nozzle cover [0110] 31 Control blade of the convergent nozzle portion [0111] 32 Control blade of the divergent nozzle portion [0112] 33 Connecting rod [0113] 34 Actuator of the convergent nozzle portion [0114] 35 Actuator of the divergent nozzle portion [0115] 36 Cooling air duct [0116] 37 Deflector plates [0117] 38 Distance [0118] 40 Outer wall