DE3710914A1 - Variable-geometry aircraft - Google Patents

Abstract

The variable-geometry aircraft has two half-wings (15a, 15b) which can be swivelled about vertical axes of rotation, as well as a horizontal tail assembly (18) and a canard empennage (16a, 16b) located forwards on the fuselage (19) and likewise capable of swivelling like the wing (15). In order to guarantee the stability about the lateral (pitch) axis (4) when the wing (15) assumes different sweeps for the purpose of adapting to different flight conditions (situations), the wing (15) and canard empennage (16) execute swivelling movements in mutually opposite directions. The respective swivel angles ( alpha , beta ) can be controlled individually or coupled synchronously depending on requirements (e.g. extension of the landing flaps (22), flight with optimum drag-to-lift ratio). Moreover, the wing (15) and canard empennage (16) can be simultaneously completely retracted into the fuselage (19), as is advantageous for various purposes. For high-speed flight, the wing (15) is retracted completely into the fuselage (19), while the canard empennage (16) guarantees the longitudinal stability by means of a specific swivel position (16a'<v>). The situation is the same in the case of intermediate positions of the wing (15) and, by relieving the horizontal tail assembly (18), this almost entirely maintains the entire trim range thereof for controlling the flight path. Reliable control of all the flight ranges from high-lift flight to cruising flight with an optimum drag-to-lift ratio up to high-speed flight is thereby rendered possible. <IMAGE>

Description

The invention relates to a so-called variable aircraft Geometry with device to adapt to different flight states, with the wing halves about vertical axes of rotation swivel and tail and tail fins on the fuselage tail are arranged.

Such aircraft in training as swivel wing aircraft and missiles are known in a number of embodiments such. B. the types Grumman F-14 (hereinafter referred to as A) and Boeing AGM-86B Cruise Missile (= B), which are described in JANE'S "All the World-s Aircraft 1985-86" and shown in Fig. 1 are.

At A the variable geometry of the main wing is used for adaptation to different flight conditions, z. B. both the maximum up drove for the landing approach as well as the best glide ratio for the Cruise is achieved by fully swinging out the wing. In contrast, there is the least resistance for the fast fly at the maximum arrow position of the wing. Intermediate positions of the wing to adapt to certain missions are included easily possible.

Category A aircraft, however, have the configuration conditional disadvantage that the pivot points of the wing halves relative on a fixed wing far away from the fuselage axis must be arranged inside. When placing the pan axles further inward would be the centers of buoyancy of the wing halves so far back when swiveling, that the aircraft no longer flies horizontally from the tail heights tail could be held, d. H. the moment balance around the transverse axis (= longitudinal stability) would be out of control. Due to the relatively short swivel wing halves at A can the lift potential of the entire wing cannot be used in this way as would be the case if the pivot points were further inside find and thus not only a part of the wing half swiveling would.

For the fast flight the wing halves are in maxi painterly arrow position, whereby it is disadvantageous here that the Wings cannot be completely swung into the fuselage, so as to minimize the resistance.  

To support the longitudinal stability during the wing pivoting process is A with small left and right from the fixed wing inside triangular stabilization surfaces equipped, which, however, in their effectiveness on the longitudinal stability because of the compact plan form and the small lever arm to the transverse axis are rather limited.

Another aircraft configuration avoids this disadvantage, where the stabilization completely separate from the wing surface (duck elevator) is arranged on the front of the fuselage, whereby a lever arm favorable for longitudinal stability results. Such a configuration represents e.g. B. the aircraft type Beech Starship I (= C, source as above) represents with a partial swiveling duck tail, which is also slimmer and therefore is more effectively trained for slow flight. It serves as a duck tail in addition to the compensation of the nickmo from the flap deflection primarily to to ensure the longitudinal stability of the duck plane since there is no further tailplane on this type of aircraft is. For this reason, the duck tail at C must constantly more or less unfolded, but what because of him caused downwind on the wing nega behind tive influence on flight performance and control. Incidentally, category C aircraft with rigid wings not at all to the extent of different flight conditions adaptable as it is with Category A aircraft Swivel wings is the case.

With the missile B mentioned at the beginning with complete swivel wings of great extension that can be folded into the fuselage (= Square of the wingspan / wing area) and tail unit The variable geometry is not fin - as suspected could - for the adaptation to different flight conditions (cf. A) provided, but it only has the task of Place wings in the fuselage to save space. To this end and to take full advantage of the wing stretch for a good one The slim wing halves have their twist point inside the fuselage, so that the fixed one at A Inner wing part omitted.  

The disadvantage of concept B, however, is that the variable Geometry cannot be used for several flight states. The The reason is that caused by the wing swing great pressure point hike, which from conventional heights tail can no longer be compensated. A stable flight condition with range performance with only partially pivoted Wings would therefore not be feasible since the aircraft has a have too much head heaviness.

The invention was therefore based on the object of an aircraft propose variable geometry, where under the full Utilizing the wing stretch to adapt to different Flight conditions are given without sacrificing longitudinal stability. In addition, with the arrangement according to the invention, the above-mentioned. Disadvantages of configurations A, B and C can be avoided and thus the area of application will be significantly expanded. In particular special should be avoided that only part of the full on drive potential of the wing is usable if the wing is completely is swung out. On the other hand, the aircraft resistance can be reduced even further at the maximum wing arrow position can as z. B. is the case with A; also different Wing pivot positions from the equilibrium of moments around the cross do not lead to an uncontrollable flight state, as would be the case with B. Thus, the inventive Aircraft good adaptability to a wide variety Have flight conditions.

In addition, no performance should be in flight with the best glide ratio loss due to the downward wind of the duck tail. Another advantage of the invention should be that no or only a small amount when the flaps are extended Downforce needs to be applied to the tail elevator. For normal aircraft or category A aircraft must go against the landing approach from the tailplane a remarkable Land performance reducing downforce are generated.

Another object to be achieved by the invention is Improve the control effectiveness of the tail elevator called in all flight conditions, which over a larger Trimming should have z. B. A.  

Starting from the aircraft of variable geo mentioned at the beginning metry with two pivoting main wing halves and tail guide factory fins there is the solution of the task in that the pivotable main wing is larger than the wing Elongation is formed and that an additional at the front Fuselage tailplane similar to the main one wing has two pivot bearings in the fuselage, which during the Swiveling the main wing through its own opposite direction running swivel movement of the duck tail the moments Ensure balance around the transverse axis.

Assuming a certain design of wing and The tail and the center of gravity should be the duck tail with the wing fully swung out, disappear into the fuselage in order to thus avoiding the harmful downwind on the wing and thus one to be able to achieve the optimal glide ratio for the cruise. In this In this case, flight control is carried out entirely by the tail unit fins and controls on the wing.

Around both duck tail and wing completely in the fuselage to be able to make disappear, are the pivot points of the pan surfaces arranged according to the fuselage.

Other flight conditions in which the full swing of the Wing is advantageous are takeoff and landing. The usual flaps provided on the wing cause by their Deflection a pitching moment that with a conventional stern arrangement of the horizontal stabilizer he downforce on the horizontal stabilizer demands. By dosing the dosing according to the invention Duck tail can not only avoid this downforce, it can even the overall buoyancy increases and the control area of the Tail unit can be improved by the fact that almost the full trim area of the tail unit even when the tail is deflected Flaps are available. It can also be a focus hike by adjusting the duck tail in addition to off blows of the tailplane can be easily compensated. Too far Additional lift can be provided on the duck tail be.  

If higher speeds are required for flight operations, is the arrow position of the wing halves for low resistance required, according to the invention the duck tail by a own opposing swiveling movement the big - because Lack of a fixed wing inner part caused - pivot point compensates for the hike.

For the flight state with minimal resistance when fully on folded main wing ensures an appropriate extension the duck tail surfaces the moment equilibrium around the cross axis. For the task dive or space-saving sub all swivel surfaces in the fuselage can be folded in.

Further features and advantages of the invention go from the sub claims and the description.

An embodiment of the invention is described below with reference to Drawings explained. It shows

Fig. 1 shows an overview of the aircraft category A, B and C,

Fig. 2 is a perspective view of the aircraft according to the invention,

Fig. 3 top view and front view of the flight device according to the invention with different configurations of swivel surface positions.

In Fig. 1-B the missile Boeing AGM-86B is shown in perspective with pivotable wing halves 1 a , 1 b , wing pivot bearings 2 , fuselage longitudinal axis 3 and transverse axis 4 . Vectors denote weight G , lift in normal wing position L and swivel position L ' and the pitching moment M' . When pivoting the wing halves 1 a , 1 b in the position shown in dashed lines moves in the wing half of the buoyant center point 5 in the position 5 ' , which causes a shift in the total drive L by the distance x to the position L' (= pressure point migration ) and the pitching moment M ' around the transverse axis 4 . In this case, the tail elevator is not sufficiently effective to compensate for this moment and the missile could no longer be caught from the resulting dive.

Although the slender wing 1 uses its lift potential well by arranging the pivot bearings 2 in the fuselage (the entire wing with high elongation lies in the air flow), it is not adaptable to other flight missions where a larger wing sweep is required.

The dash-dotted wing position 7 shows the fully collapsed state in the fuselage, which is advantageous for trans port on the ground and in the air and for accommodation.

Another aircraft with variable sweeping embodies the aircraft Grumman F-14 shown in Fig. 1-A, in which the wing halves 8 a , 8 b pivot about the pivot bearing 9 , which are on the outside of the fixed wing parts 10 . With this arrangement, however, only a part of the total wing area can pivot, which means that the full lift potential (as in B) cannot be used. In order to ensure stability around the transverse axis in high-speed flight at the maximum wing arrow position, pivotable stabilizing surfaces 11 are provided from the inner wing part 10 . 11 ' shows the pivoted-in state. The effectiveness of these stabilization surfaces 11 is rather limited due to the compact fit and the low lever arm.

The maximum wing arrow position 8 a ' , 8 b' shows that the wing halves cannot be pivoted completely into the fuselage and it is therefore not possible with configuration A to further minimize the resistance for high-speed flight.

In the Fig. 1-C type Beech Starship I is a duck plane without variable geometry of the main wing, which has a swiveling duck tail 12 to compensate for the pitching moment from flaps, which in landing configuration has the front position 12 ' ( dash dotted). Since there is no tail unit here, the duck unit must both compensate for the flap deflection and ensure flight control around the transverse axis. Due to this design, the duck tail is subject to high aerodynamic loads, which in particular leads to swirling pegs 13 (drawn only on one side) on landing, which disrupt the wing flow and cause considerable control problems; especially with cross winds. In addition, there is a restriction on the height control, since a large part of the available trim area of the duck tail 12 must be used for the compensation of the flap deflection. Otherwise, a wide adaptation to different flight conditions is not given in configuration C, since here the wing 14 is permanently installed and has no pivoting device.

In Fig. 2, the aircraft according to the invention is shown in perspective in the design of an aircraft with a high Mach number, the pivotable wing halves 15 being completely folded into the fuselage while the duck tail 16 is partially pivoted out about the transverse axis 4 to compensate for the top-heavy moment ( 16 a ^IV ). The dash-dotted wing halves 15 a ' , 15 b' and duck tail halves 16 a ' , 16 b' show the respective maximum pivoting position. At the rear of the aircraft there are tail fins for the side control 17 and height control 18 , which can also be pulled for roll control about the fuselage longitudinal axis 3 (Taileron).

Incidentally, it should be mentioned that the invention is not only limited to the exemplary embodiment in FIG. 2, but other aircraft with other uses can also be considered.

Such an example, in which it also depends, mini paint resistance in fast flight and maximum lift in Ensuring slow flight would be a space shuttle. A other applications would be unmanned missiles. In case of a Space shuttle can use the proposed variable geometry too one compared to today's usual ferries with delta wings significantly improved glide ratio, leading to safety Significantly expand the scope when selecting landing sites could, since normal major airports were now sufficient. Furthermore would the ferry according to the invention over a today Ferries have a greatly reduced surface, which is particularly important when it comes to Re-entering the earth's atmosphere by reducing the Insulation materials have a favorable effect on the payload ver relationship.

Fig. 3 shows a plan view and a front view of the aircraft according to the inven tion. Different wing and duck tail positions are shown on the left and right in the top view, which can be assigned to different flight conditions. The halves of the swivel wing 15 are both in the fully pivoted position 15 a ' , 15 b' rotatable about the pivot bearing 2 and fully foldable into the fuselage 19 (position 15 '' ). The rotational position of the duck tail halves, which is also arranged in the fuselage, is labeled 21

When flying with the best glide ratio, the wing is fully swung out into the position 15 a ' , 15 b' while the duck tail fully occupies the position 16 a ' , 16 b' . In this flight state, the mutual swivel control of wing and duck tail can be canceled arbitrarily.

For fast cruising, the wing halves pivot from the maximum swivel position 15 a ' , 15 b' to an intermediate position 15 a ''' , 15 b''' , the swiveled angle is designated by α . To compensate for the resulting pressure point shift to the rear, the duck tail 16, starting from the pivoted position 16 a ' , 16 b', runs in the opposite direction to the wing 15 pivoting movement into the position 16 a ''' , 16 b''' . The angle from here to the maximum pivoted position 16 a '' , 16 b '' is denoted by β , which, depending on the flight mechanical requirements, can be controlled individually or can be coupled with the pivoting angle α of the wing 15 syn chronically.

By pivoting the duck tail 16 , the moment about the transverse axis 4 can be substantially compensated for, so that the flight path is controlled by the tail tail fins 17, 18 for this flight condition, for which purpose the full trim area is available. For the control of the respective pivoting angles α , β and their assignment to the flaps from the flaps 22 and possibly existing ducks, the tailplane 23 should advantageously be provided with an electronic control system.

As shown in Fig. 1B, when swinging a slim wing occurs the pressure point migration x with the resulting moment M ' . In the case of the wing pivot position 15 a ''' shown in Fig. 3, the pressure point migration x''' (on the drive center 5 ' moves to 5''' ) by the position of the duck tail 16 a ''' compensated by the pro duct of buoyancy in point 20 ''' and lever arm y''' . The description only refers to one side of the aircraft.

Are the flaps 22 extended (shown on one side) with the wing fully pivoted in the position 15 a ' , 15 b' , so the pressure point 5 ' to 5'' moves by the distance x'' to the rear, which is a top-heavy moment causes that can be counteracted by a dosed swiveling of the duck elevator in position 16 b '' ent (buoyancy in 20 '' and lever arm y '') . In this case, too, the opposite control of wing 15 and duck tail 16 can be decoupled and switchable to superimposed arbitrary control.

It should be noted here that because of the large lever arm y '' , the duck tail only needs to have a small lift coefficient. As a result, the resulting weak vortex drag cannot impair flight performance.

If minimal flight resistance is required in high-speed flight, the wing halves can be completely folded into the fuselage 19 and take the position 15 '' . To compensate for the resulting top heaviness, which is however reduced by the buoyancy of the fuselage, the duck tail swings into position 16 a ^IV .

To fulfill special purposes (transport on the ground and in the air, storage) both wing and duck tail (position 15 '' and 16 a ' , 16 b') can be completely folded into the fuselage 19 .

Claims (7)

1. aircraft variable geometry with means for adapting to different flight conditions, the wing halves are pivotable about ver tical axes of rotation and vertical and vertical fins are arranged on the fuselage tail, characterized in that the pivotable main wing ( 15 ) is designed as a large extension wing and that an additional front of the fuselage ( 19 ) located duck elevator ( 16 ) similar to the main wing ( 15 ) has two pivot bearings ( 2 ) or ( 21 ) in the fuselage ( 19 ), which during the pivoting of the main wing ( 15 ) by a own counter-rotating swivel movement of the duck tail ( 16 ) ensure the moment balance around the transverse axis ( 4 ). 2. Aircraft according to claim 1, characterized in that the respective pivot angle ( α or β ) of wing ( 15 ) and duck elevator ( 16 ) can be individually controlled and / or synchronously coupled. 3. Aircraft according to claim 1 and 2, characterized in that both the main wing ( 15 ) and duck tail ( 16 ) can be completely folded into the fuselage ( 19 ). 4. Aircraft according to claims 1-3, characterized in that in order to achieve the best glide ratio the wing is fully pivoted ( 15 ' ), while the duck tail is completely folded into the fuselage ( 19 ) ( 16' ). 5. Aircraft according to one of claims 1-4, characterized in that the full pivoting of the wing ( 15 ) in a fixed pivot position ( 15 ' ) the opposite pivot control of the wing ( 15 ) and duck tail ( 16 ) can be canceled arbitrarily. 6. Aircraft according to one of claims 1-5, characterized in that in the area of slow flight with extended flaps ( 22 ) and fixed wing position ( 15 ' ), the Ge guarantee of the equilibrium of moments about the transverse axis ( 4 ) by metered pivoting of the duck tail ( 16 ) and the opposite control can be decoupled and switched to a superimposed arbitrary control. 7. Aircraft according to claims 1-6, characterized in that in the about the transverse axis ( 4 ) - due to a corresponding pivoting position of the duck elevator ( 16 ) - balanced flight condition, the control of the flight path is taken over by the tail fins ( 17, 18 ).