DE102008022452A1 - Aircraft, has central flight controller adapted such that individual auxiliary wings are adjusted in position independent of other auxiliary wings, where position of auxiliary wings is adjusted to each other and to main wings - Google Patents

Abstract

The aircraft has left and right auxiliary wings (2L, 2R) movably and actively arranged at an outer end of left and right main wings (1L, 1R). Left and right outer auxiliary wings (3L, 3R) are connected to the auxiliary wings in a wing span direction of the main wings. A position of the auxiliary wings is adjusted to each other and to the main wings in degrees of freedom of movement at axes running parallel to an aircraft longitudinal axis (Xf). A central flight controller (F) is adapted such that individual auxiliary wings are adjusted in a position independent of the other auxiliary wings.

Description

The The invention relates to an aircraft with swept main wings, At the outer end of each auxiliary wing are movably and actively arranged controllable.

Aircraft of this type, in particular those with a central hull, are known from the US-PS 59 88 563 , from the US 2004/00619 , from the DE 324 25 84 , from the DE 103 12 290 and from the DE 10 2005 028 688 ,

In the aforementioned documents is disclosed, the movable auxiliary wings for influencing the buoyancy and impact on the aircraft To use resistance.

From the Publication AIAA 2006-3660 (Forth Applied Aerodynamics Conference - June 8, 2006, San Francisco, California) entitled "The application of variable cant angle winglets for morphing aircraft control" It is known, auxiliary wings, which are pivotally mounted at the outer end of the main wing and are independently adjustable in their inclination to the main wing to use as a means for the roll yaw and pitch control of an aircraft. According to 14 the above-mentioned document, in addition to a first pair of auxiliary wings, a second pair of such auxiliary wings are used, wherein the arrangement disclosed in the aforementioned document provides the two auxiliary wings at each end of the main wing in the direction of flight one behind the other. With the arrangement according to the above-mentioned AIAA paper, control forces can be exerted on an aircraft, which enable the curved flight by coupled rolling and yawing and insofar as at least partially replace conventional control means for an aircraft. An exercise of control functions that control an aircraft independently of each other with respect to the three major axes of the aircraft, namely longitudinal axis, transverse axis and vertical axis (corresponding to roll control, pitch control and yaw control) is not possible with the arrangement disclosed therein.

task the present invention is by means of independently controllable auxiliary wing the entire flight control one Aircraft in a controlled manner, thereby reducing conventional control including elevators, rudders and ailerons corresponding tail units at least partially, but preferably completely replace.

According to the invention the task is solved by having an airplane with the features of the preamble of claim 1 to a first auxiliary wing in the spanwise direction of the main wing a second outer auxiliary wing connected is and the position of the auxiliary wings to each other and to Main wing adjustable in given degrees of freedom of movement is, with the associated control being adapted in such a way that every single auxiliary wing is independent of the other auxiliary wings in position adjustable is.

By the inventive arrangement can aerodynamic forces and corresponding moments around the Main axes of the aircraft can be applied exactly controlled and it can thus provide a complete primary flight control, a trim around the longitudinal axis of the aircraft and high-lift and an air brake can be realized. Here is the adjustability the second (outer) auxiliary wing for the achievement of independent motion quantities of the aircraft with respect to the three main aircraft axes of special meaning.

The inventive principle would also be realized if in addition to the two adjoining auxiliary wings to the outside even more controllable auxiliary wings would be provided.

Based on a schematic representation of an aircraft main wing with two adjoining auxiliary wings after 7 The attached drawings, the principle of operation of an aircraft according to the invention will be explained first, before embodiments of the invention will be described with reference to the further accompanying drawings.

In the schematic representation after 7 is a wing configuration with a main wing 1 , a first auxiliary wing 2 and a second outer auxiliary wing 3 shown. The main wing 1 is swept towards the aircraft longitudinal _axis x _F by an angle β to the rear. A sweep in the opposite direction forward would also be conceivable; on the other hand, the invention could not have a main wing extending perpendicularly to the main axis of the aircraft x _F 1 be achieved. The first auxiliary wing 2 should be pivotable about the axis x ₁ , which is aligned substantially parallel to the aircraft longitudinal axis x _F. The second, outer auxiliary wing 3 is about an axis x ₂ , which is substantially parallel to the axis x ₁ , pivotally mounted on the first auxiliary wing 2 be arranged. In addition, the first auxiliary wing 2 about an axis a ₂ and the second outer auxiliary wing 3 be pivotable about an axis a ₃ . The inflow direction of the wing configuration is indicated by the arrow U, which is parallel to the aircraft longitudinal axis x _F. The three main axes of the aircraft are the longitudinal axis x _F the transverse axis y _F and the vertical axis z _F. The center of gravity of the aircraft is plotted with CG, while the aerodynamic center, which is the point of action of the total of aerodynamic forces on the aircraft with AZ be is drawing. The aerodynamic center AZ is followed by dashed arrows indicating a displacement of the aerodynamic center relative to the stationary center of gravity CG depending on the control of the auxiliary wings 2 and 3 to indicate. The rotation arrows around the x _F , y _F and z _F direction indicate roll moment, pitch moment and yaw moment in this order.

in the stabilized flight (trimmed flight condition) agrees the location of the AZ always match the location of the CG.

As can be understood from the schematic illustration, the total of the aerodynamic forces by size and the point AZ changes by changing the position of the auxiliary wings 2 and 3 as follows:
By pivoting the auxiliary wing 2 about the axis x ₁ the force component changes parallel to the y _F -axis (= A ₂ y) and the force component A ₂ z parallel to the z _F -axis (= A ₂ z). The further the auxiliary wing 2 is pivoted out of the x _F y _F plane, the greater becomes A ₂ y at the expense of A ₂ z. By symmetrical pivoting of the auxiliary wings 2 on both sides of an aircraft (in the schematic illustration 7 if only one side of the aircraft is shown), the amount of resulting total lift will decrease from the x _F -y _F plane. At the same time, the position of the aerodynamic center AZ shifts in the positive x _F direction, whereby a moment about the y _F axis is exerted on the aircraft, ie a positive pitching moment. When swinging the auxiliary wings 2 into the x _F y _F plane the force component A ₂ z increases. At the same time, the position of the aerodynamic center AZ moves in the negative x _F direction. Thus, a larger lever arm is created between the resulting total lift vector and the center of gravity of the aircraft. The result is a negative pitching moment (aircraft nose down). The same direction pivoting two symmetrically to the aircraft longitudinal axis x _F arranged auxiliary wing 2 can be used as a means of high lift (secondary flight control) due to the resulting variation in the total lift vector.

The simultaneously generated pitching moment about the y _F -axis can be used as a means for trimming the aircraft in the longitudinal axis or facilitate trimming.

A movement of the auxiliary wings 2 asymmetrical to the x _F z _F plane, ie different on both sides of the aircraft, leads to a change in the buoyancy components A ₂ z. This is done when panning from the x _F y _F plane according to 7 primarily a shift of the aerodynamic center AZ in the positive y _F direction. Due to the thus generated lever between the shifted in the y _F direction position of the AZ and the x _F axis results in a negative rolling moment about the x _F axis. In addition, however, this also leads to a shift of the AZ in the positive x _F direction, resulting in a positive pitching moment. At the same time, the force component A ₂ y acts in such a way that due to the distance between the z _F axis and the point of application of the force component A ₂ y, a curve-promoting yaw moment is produced about the z _F axis.

By the to the auxiliary wing 2 hinged auxiliary wing 3 can one of the auxiliary wing 2 realize decoupled yaw and pitch control of an aircraft equipped in this way. By superposition of the function of the auxiliary wing 3 with the function of the auxiliary wing 2 on both sides of the aircraft a decoupled yaw, pitch and roll control, air brake and high lift can be realized.

Based on a trimmed flight condition of a suitably equipped aircraft take the auxiliary wings 2 . 3 an initial angle around the a ₂ -axis and a ₃ -axis. Due to the flow U and a resulting edge vortex is on the auxiliary wing 2 and 3 generates a buoyancy vector with components in the positive x _F direction and y _F direction. In that regard, results in a conventional winglet functionality. By the same direction deflection of the symmetrically arranged to the x _F z _F plane auxiliary wing to the a ₂ -axis and / or a ₃ -axis, the function of an air brake can be realized. Due to the great distance of the auxiliary wings 2 and or 3 From the fuselage a vertical stabilizer function is exerted on the aircraft, which can replace a conventional rudder.

Is by pivoting the auxiliary wing 2 A curve maneuver is introduced about the axis x ₁ differently on both sides of the aircraft, so that the respective rolling, yawing and pitching movements of the aircraft can be achieved by pivoting the auxiliary wing 3 be amplified or damped around the x ₂ axis and / or around the a ₃ axis. For this purpose, by means of the isolated control by the auxiliary wing 3 generated aerodynamic moments with that of the auxiliary wing 2 superimposed moments generated. Here is the direction and the size of the pivoting of the auxiliary wing 3 around the x ₂ -axis and the a ₃ -axis crucial for the expression and orientation of the buoyancy force vector A ₃ and thus for the size and direction of the superimposed moments. If z. B. the auxiliary wing 3 pivoted with its leading edge to the outside, so there is a resulting lift vector, which is also directed to the outside and it is generated a positive yaw moment about the z _f -axis of the aircraft. An increase in the effect is, as can be easily seen, achieved in that on the left and right wing of an aircraft, the auxiliary wing 3 is pivoted asymmetrically to the x _F z _F plane. In addition, through different setting of the leading edge of the auxiliary wing 2 with respect to the a ₂ -axis, which influences roll, yaw and pitch movements.

A advantageous embodiment, the invention undergoes thereby, that in addition to the adjustment of the auxiliary wing in several degrees of freedom control of the profile geometry the auxiliary wing provided in the sense of a change in the lift coefficient is. The profile geometry of the auxiliary wing can be in the way changed that a shape change in the sense of a change in profile curvature and Thickness takes place, resulting in the lift coefficient of a wing profile known to change. Another change of Profile geometry in the sense of a change in the lift coefficient can be done by the auxiliary wing over its span length is twisted thereby over the Spanlängelänge a different angle of attack and thus again results in different buoyancy. The change the profile geometry can be done in a continuing education, that they have a continuous curvature change from positive to negative profile curvature, whereby then the direction of the lift vector is also reversed can be.

A further advantageous embodiment undergoes the invention in that an actuating device for movement the first auxiliary wing is adapted so that they themselves adjoining outer auxiliary wing as an actuator. With such an actuator a movement of the first auxiliary wing so in the Way, that first about the central control and a separate actuator the outer Auxiliary wings are adjusted so that their force or Moment on the first auxiliary wing, where they struck are, in the sense of the desired adjustment of the first auxiliary wing acts. So it will be the resulting from the aircraft approach aerodynamic forces and moments of the outer Auxiliary wing as an actuator for the first auxiliary wings used. By combination with a suitable locking device can a change of shape once with the help of Aktuatorkräfte be permanently obtained from the auxiliary wings.

Thereby can an otherwise necessary actuator device for the adjustment of the first auxiliary wing (z. As hydraulic servomotors or electric drives) accordingly be trained weaker or you can on such additional work actuators for adjusting the first Auxiliary wings completely dispense, resulting in the With regard to construction costs and construction weight very beneficial.

Finally, according to a preferred embodiment of the invention, the setting of the auxiliary wings and / or the change in their profile geometry by structural deformation. Such a design has the advantage that corresponding machine elements, such as bearings, including the associated lever mechanism are largely dispensable. It then only has the structure at the points of the desired deformation, so z. B. along the pivot axes x ₁ , x ₂ or a ₂ and a _3, the structure of the wings or auxiliary wings have an actively adjustable stiffness. In addition, corresponding means for blocking the deformation structure in the desired end position may be provided. In this case, in a further embodiment of the invention, a structural deformation of the main wing with the help of Aktuatorkräfte or moments of the auxiliary wings may be provided.

Based The accompanying drawings will be an embodiment of the invention explained with different configurations.

In the drawings shows

1 the oblique image of an aircraft with swept main wing and auxiliary wings arranged thereon and an associated control schematically,

2 an enlarged view of the outer wing portion of the main wing 1L out 1 with marked swivel and pitch axes and angles,

3 - 5 Representations of the outer wing section 1L according to 2 with differently set auxiliary wings 2 . 3 .

6 a simplified image of an aircraft according to the invention in cross section in eight different flight conditions depending on the setting of the auxiliary wings and

7 a schematic representation of an aircraft main wing with two auxiliary wings arranged thereon with functional sizes.

This in 1 schematically shown in an oblique plane has a fuselage 10 and arranged rear winged main wing 1L (left) and 1R (right). The sweep angle with respect to the aircraft longitudinal axis x _F is denoted by β, the wing leading edge with 12 , At the farthest end of the main wing 1L . 1R is pivotable about a first pivot axis x _1, a first auxiliary wing 2L respectively. 2R on the main wing 1L respectively. 1R (right) struck. At the outer (fuselage) end of the first auxiliary wing 2L respectively. 2R is in each case a second (outer) auxiliary wing 3L respectively. 3R pivotally articulated about an axis x ₂ (L, R). The two auxiliary wings 2 . 3 are the other spanwise aligned axes a ₂ and a ₃ pivotable. The operation and control of the auxiliary wings 2 . 3 takes place actively by the pilot (P) via a central flight control F, from where the size and direction of the pivoting movement of the auxiliary wing 2 . 3 be set by the pilot F according to the pilot. The sizes to be set are 1 the angle α ₂ , the angle α ₃ , and the angle φ ₂ , φ ₃ specified. These angles are in the illustration below 2 indicates, from which it follows that the angle α _{2 of} the inclination angle between the first auxiliary wing 2 and the main wing 1 and α _{3 is} the corresponding inclination angle of the second, outer auxiliary wing 3 opposite the auxiliary wing 2 is. With φ2 and φ3, the tilt angles are the auxiliary vanes 2 . 3 designated by the axes a ₂ and a ₃ . The flow of the aircraft according to 1 or the wing section according to 2 takes place in the direction of arrow u. The practically technical training of the pivoting of the auxiliary wing 2 . 3 about the pivot axes x ₁ , x ₂ and a ₂ , a ₃ is not shown in detail. It may be conventional swivel joints with corresponding actuators, as are well known in the art. However, it can also, as indicated in the figures, in a preferred embodiment, a pivoting about the said pivot axes be provided in such a way that the auxiliary wings 2 . 3 integral to the main wing 1 are formed, wherein along the desired pivot axes x ₁ , x ₂ , a ₂ , a _{3 is provided} an actively adjustable stiffness of the wing structure. The pivoting of the auxiliary wings is then carried out with the angles α ₂ , α ₃ and φ ₂ , φ ₃ by appropriate means for changing the shape.

At the in 2 The illustration shown is the outer auxiliary wing 3 detached from the first auxiliary wing 2 shown and only the first auxiliary wing 2 integral to the main wing 1 molded and pivotable by means of deformation about the pivot axes x ₁ and a ₂ . A pivoting of the outer auxiliary wing 3 about the axis a ₃ should be done by means of a conventional pivot joint, not shown here. Instead of a pivoting of the outer auxiliary wing 3 around the axis a ₃ can also change the profile of the outer auxiliary wing 3 be provided in the sense of a positive or negative profile curvature or a change in the profile thickness to the buoyancy value of this outer auxiliary wing 3 to vary.

In the 3 to 5 Again, the outer wing area is analogous to the representation in FIG 2 shown, wherein here an embodiment is shown, in which the adjustment of the two auxiliary wings 2L . 3L is effected both with respect to the angle α ₂ and α ₃ as well as with respect to the angle φ ₂ and φ ₃ exclusively by changing the shape of the wing structure. The representations after the 3 . 4 and 5 differ only by different settings of the auxiliary wings 2L . 3L in terms of inclination and angle of attack.

In 5 is at the main wing 1L another area 4 marked within the joint area 5 between the first auxiliary wing 2L and the main wing 1L lies. By targeted variation of the structural rigidity of the main wing 1 in this area or in the area 5 can be achieved by aerodynamic forces from the auxiliary wings 2 and 3 a deformation in the sense of a buoyancy change of the main wing 1 he follows. In this case, by means of suitable locking devices, once a change in shape with the help of Aktuatorkräfte from the auxiliary wings 2 . 3 be permanently obtained until the locking device is released again without causing a permanent force by the auxiliary wing 2 . 3 would be required.

The representation after 6 is self-explanatory and should only indicate schematically, which adjustment of the auxiliary wing leads to which flight condition. The right lower illustration "additional degrees of freedom" shows a flight situation in which the yaw rate of the aircraft can be changed simultaneously during the rolling process.

1

Aircraft main wing

2

first auxiliary wing

3

second (outer) auxiliary wing

U

flow direction

x _F

aircraft longitudinal axis

y _F

Plane transverse axis

z _F

Aircraft vertical axis

CG

Airplane CG

AZ

aerodynamic Center (point of action of the total of aerodynamic forces)

A1

Aerodynamic force (buoyancy) of the main wing 1

A2

Aerodynamic force (lift) of the auxiliary wing 2

A3

Aerodynamic Force (buoyancy) of the outer auxiliary wing

x ₁

Swivel axis of the auxiliary wing 2 on the main wing 1

x ₂

Swivel axis of the second auxiliary filler 3 on the first auxiliary wing 2

a ₂

swivel axis perpendicular to x1

a ₃

swivel axis perpendicular to x2

α ₂

Tilt angle between auxiliary wing 2 and main wing 1

α ₃

Tilt angle between auxiliary wing 3 and auxiliary wings 2

ß

V-angle the main wing

φ ₂

swivel angle around the axis a2

φ ₃

swivel angle around the axis a3

QUOTES INCLUDE IN THE DESCRIPTION

This list The documents listed by the applicant have been automated generated and is solely for better information recorded by the reader. The list is not part of the German Patent or utility model application. The DPMA takes over no liability for any errors or omissions.

Cited patent literature

• US 5988563 [0002]
• US 2004/00619 [0002]
• - DE 3242584 [0002]
• - DE 10312290 [0002]
• - DE 102005028688 [0002]

Cited non-patent literature

• Document AIAA 2006-3660 (Forth Applied Aerodynamics Conference - June 8, 2006, San Francisco, California) entitled "The application of variable cant angle winglets for morphing aircraft control" [0004]

Claims (8)

Airplane with swept main wings at the outer (fuselage) end of each auxiliary wings are arranged movable and actively controllable, characterized in that on a first auxiliary wing ( 2R . 2L ) in the spanwise direction of the main wing ( 1R . 1L ) a second outer auxiliary wing ( 3R . 3L ) and the position of the auxiliary wings ( 2R . 2L ; 3R . 3L ) is adjustable to each other and to the main wing in predetermined degrees of freedom of movement, wherein the associated control is adapted such that each individual auxiliary wing ( 2R . 2L ; 3R . 3L ) is adjustable in position independently of the other auxiliary wings. Aircraft according to claim 1, characterized in that a first degree of freedom of movement for the auxiliary wings ( 2R . 2L . 3R . 3L ) is a pivoting about substantially parallel to the aircraft longitudinal axis (x _F ) extending axes (x ₁ , x ₂ ). Aircraft according to claim 2, characterized in that as additional movement degree of freedom for the auxiliary wings ( 2R . 2L . 3R . 3L ) A pivoting about a transverse to the auxiliary blade profiles extending axis (a ₂ and a ₃ ) is provided (change of the angle of attack φ ₂ , φ ₃ ). Aircraft according to claims 1 to 3, characterized in that by means of the control the profile geometry of the auxiliary wings ( 2R , L, 3R , L) is variable in terms of a change in the lift coefficient. Aircraft according to claims 1 to 4, characterized in that an actuating device for moving the first auxiliary wings ( 2R . 2L ) is adapted so that the adjoining outer auxiliary wing ( 3R . 3L ) as an actuator. Aircraft according to claims 1 to 5, characterized in that the adjustment of the auxiliary wings and / or the change of the profile geometry by structural deformation he follows. Aircraft according to claims 1 to 6, characterized in that the size of the auxiliary wings and their range of motion are so dimensioned and tuned to the stiffness of the main wing, that a usable for the flight control change of the buoyancy coefficient of the main wing by its change in shape in at least one hull lying area ( 4 ) is achievable. Aircraft according to claims 1 to 7, characterized in that the auxiliary wings are integrally formed on the main wing and the wing structure in the areas ( 5 . 6 ) has the desired deformation by controlled movement of the auxiliary wings an actively adjustable stiffness.