JP2005088804A - Variable delta wing airplane and its fuselage attitude controlling method - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a variable delta wing airplane equipped with a delta wing receiving a less air resistance during supersonic flight. <P>SOLUTION: The variable delta wing airplane embodied in a lightweight construction and receiving a less air resistance is structured so that the delta wing 3 as the main wing in a triangular shape fundamentally is formed from an inner wing 4 and an outer wing 5, in which the purpose of the invention is accomplished by rotating the outer wing 5 round the Mach line 7 symmetrical to the left and right which is aslant to the fuselage axis. <P>COPYRIGHT: (C)2005,JPO&NCIPI

Description

  The present invention relates to a high-speed aircraft having a delta wing structure, and more particularly, to a high-speed aircraft having a variable delta wing structure having an improved delta wing and a method for controlling the attitude of the aircraft.

  Traditionally, in high speed aircraft flying at supersonic speeds, delta wings are often used to reduce air resistance at supersonic speeds. Delta wings have a lower lift coefficient than straight wings, so the lift tends to be insufficient at takeoff. For this reason, various proposals have been made so far to supplement lift during takeoff. For example, as a second wing between the main wing and the nose, a canard is installed so that it can rotate, and when taking off, the canard is deployed at right angles to the axle to compensate for lift and rotate according to the aircraft speed to optimize flight performance. There has been proposed a method in which a super cruising cruise is carried out (see Patent Document 1). In addition, a wing structure that spreads the wings to the left and right during takeoff has been proposed as a variable swept angle wing. However, in order to realize this mechanism, it is necessary to increase the strength of the airframe structure, which increases the weight of the airframe. In addition, there is a problem that the volume for mounting the fuel in the blade is insufficient. Furthermore, since the cross-sectional area distribution and lift distribution are not symmetric in the front-rear direction, the resistance increases during supersonic flight, and the lift center moves backward, so the balance in the front-rear direction is lost. There is also.

The North American XB-70 aircraft has a mechanism that folds the outer wing of the delta wing downward, but the folding line is parallel to the fuselage axis. Not intended to smooth the lift distribution. Further, Patent Document 2 shows a delta wing structure having a fin steering surface at the wing tip. However, the delta wing structure is used at a low speed such as during take-off where the rudder enters the shadow of the delta wing and the steering effect becomes poor. It is not intended to reduce the movement of the center of lift during supersonic flight.
US Pat. No. 5,992,796 Japanese Patent Publication No. 6-2480

  As described above, the delta wing is adopted as a high-speed aircraft for supersonic flight. Because of the movement, there is a problem that the balance in the front-rear direction is lost and the fuel when rising is high. To make up for these drawbacks, various proposals have been made as described above. However, each has its advantages and disadvantages, and it has not been realized that has satisfactory performance in any flight phase of takeoff, subsonic flight, and supersonic flight. .

  Therefore, the present invention has been made in view of the problems of the prior art, and solves the problem of the movement of the lift center of the delta wing and the air resistance, reduces the movement of the lift center at the transonic speed, Provided is a variable delta wing structure aircraft capable of reducing air resistance during supersonic flight without sacrificing subsonic performance, saving ascending fuel, and reducing influence on the ozone layer, and its attitude control method. For the purpose.

  The variable delta wing aircraft of the present invention that solves the above problems is a main wing that is basically a triangular plane, and is composed of an inner wing and an outer wing, and the outer wing is symmetrical with respect to the axis of symmetry. It is provided to be rotatable with respect to the inner wing. It is desirable that the mounting angles of the left and right outer wings can be varied individually. A fuel tank can be provided in the inner wing.

  In addition, the attitude control method for a variable delta wing aircraft according to the present invention comprises a delta wing as a main wing composed of an inner wing and an outer wing, and the outer wing is an axis of symmetry with respect to the inner wing. The outer wing mounting angle is set to 0 degrees during take-off, and the left and right outer wings are folded upward during high subsonic or supersonic flight to prevent air resistance and change in body posture. On the other hand, the body posture is controlled by individually controlling the mounting angles of the left and right outer wings.

  According to the present invention, the supersonic resistance of the delta wing is reduced, the lift center movement in the front-rear direction at the time of sonic breakthrough is reduced, the flight stability is improved, the take-off performance and the subsonic performance are not sacrificed, In addition, the rising fuel can be saved and the impact on the ozone layer can be reduced. Also, by controlling the mounting angle of the left and right outer wings, the attitude of the aircraft can be controlled, the vertical stabilizer can be eliminated, and the weight of the aircraft can be reduced.

  Hereinafter, the present invention will be described based on an embodiment with reference to the accompanying drawings, but the present invention is not limited to the present embodiment. Although this embodiment shows a case where the present invention is applied to a supersonic aircraft, the present invention can also be applied to a subsonic aircraft or the like.

  FIG. 1 is a conceptual plan view showing a delta wing aircraft according to an embodiment of the present invention. In the delta wing aircraft 1 of the present embodiment, a delta wing 3 that is a main wing for levitating the fuselage 2 is composed of an inner wing 4 and an outer wing 5, and is based on a triangular plane airfoil. In the present embodiment, the outer wing 5 is attached to the inner wing 4 at a position where the attachment axis 6 is equal to the Mach line 7 and is swingable up and down about the Mach line. Arbitrary means can be adopted as the means for attaching the outer wing 5 to the inner wing 4 and the swing drive means, and the means is not particularly limited. For example, the outer wing 5 is a hinge that can rotate at a tangent to the inner wing 4. And can be controlled so as to be set at an arbitrary mounting angle by a driving device such as a hydraulic cylinder or an electric motor.

The variable delta wing aircraft of the present embodiment configured as described above functions as follows.
At the time of takeoff, the outer wing 5 is positioned on the extension of the inner wing 4 and is set in an integrated state (mounting angle 0 degree). In this case, the inner wing 4 and the outer wing 5 generate lift for levitation of the airframe 2, and the outer wing 5 is integrated with the inner wing 4 to generate lift. At that time, the center of lift exists in the vicinity of 25% of the average chord length.

  When the flight speed exceeds the speed of sound, the flow around the main wing changes, and if the outer wing 5 is integrated with the inner wing 4 and attached, the lift center moves to the rear of the aircraft by 50% of the average chord length. At the time of subsonic and supersonic flight, as shown in FIG. 2 (b), the outer wing 5 is bent upward and the mounting angle with respect to the inner wing 4 is increased to prevent an increase in air resistance. The lift of is reduced. As a result, the inner wing 4 has a lift distribution close to the front-rear symmetry, and a low-resistance body according to Jones's theory is realized. At this time, if the attachment line of the inner and outer wings is inside the Mach line, the influence of the outer wing 5 does not reach the inner wing 4. Therefore, in this embodiment, although the attachment axis line is made to correspond to the Mach angle, it is not necessarily limited thereto, and the attachment axis line may be attached to the inside of the Mach line.

FIG. 2 is an example showing a comparison between front views of a conventional delta wing during supersonic cruising and a body with the outer wing raised in the present invention.
Comparing the two from the front of the airplane, the air receiving surface of the conventional delta wing 10 is triangular, whereas the variable delta wing 3 of the present invention is close to a diamond. As a result, the conventional delta blade 10 has to support the lift load in the blade span direction near the trailing edge where the blade thickness is thin, which causes bending deformation and increases the structural weight. On the other hand, in the variable delta wing 3 of the present invention, the lift is mainly handled by the inner wing fixed to the fuselage, and the deflection can be reduced.

FIG. 3 is a front conceptual view showing a state in which individual mounting angles are changed on the left and right outer wings 5.
By separately controlling the left and right outer wings 5 and changing the mounting angle around the hinge, it is possible to control the lift generated by the outer wings. Each stability can be taken.

FIG. 4 is a diagram showing the concept of the all-axis direction lift distribution by the variable delta wing according to the present embodiment.
In the diagram, a is the overall lift distribution of the variable delta wing of this embodiment, b is the overall lift distribution of the conventional delta wing, and c is a conceptual diagram showing the lift distribution of the fuselage. . As shown in the figure, in the case of the conventional delta wing, the lift distribution is asymmetric in the front-rear direction, and the center of lift is moved backward, so the balance in the front-rear direction is lost. On the other hand, in the case of the variable delta wing of the present embodiment, it is almost symmetrical in the longitudinal direction of the axle, and the problems of movement of the lift center and air resistance can be solved. In particular, in order to reduce the air resistance experienced by the aircraft during supersonic flight, it is necessary to smooth the cross-sectional area and lift distribution of all aircraft according to the idea of the area rule. Meets that requirement. On the other hand, in the combination of the elongated fuselage and the conventional delta-type main wing, as shown in the diagram b in FIG. 4, the lift of the main wing portion rapidly decreases at the trailing edge.

  On the other hand, according to Jones et al.'S theory, since the resistance at supersonic speed is equal to the case where the nose and tail are reversed, this delta wing produces a large wave resistance. In contrast, the variable delta wing of the present invention reduces the lift of the outer wing to increase the lift at the center of the fuselage, smoothly reducing the lift in the nose and tail directions, and reducing the air resistance. .

In a normal aircraft, the attitude angle of the aircraft, that is, the pitch angle, the yaw angle, and the roll angle are controlled by an elevator, a ladder, and an aileron, respectively. On the other hand, in the present invention, the change in the mounting angle of the left and right outer wings works in the same direction at the pitch angle and in the opposite direction at the yaw angle and roll angle. Therefore, for example, as shown in the block diagram in FIG. 5, by constructing the airframe attitude angle control system, the left and right outer wing actuators can be operated independently, and the attitude can be controlled. Can be abolished.
In the block diagram of the figure, reference numeral 11 denotes a controller as a central processing unit, which is connected to a storage device 12 that can update the interference coefficient table as needed, and is connected to an actuator 13 that is a swing drive device for the left and right outer blades 5. A control signal is output, whereby the mounting angles of the left and right outer wings 5 change, and as a result, the pitch angle, roll angle and yaw angle of the aircraft 14 that change as needed are detected by the respective attitude sensors 15, which are detected by the controller 11. It is the structure which feeds back to.

  In the airframe attitude angle control system configured as described above, the controller 11 that has received the attitude angle signal in response to a command input stores in the storage device 12 the interference coefficients for the pitch angle, yaw angle, and roll angle depending on the mounting angle of the left and right outer wings 5. The control values for the pitch angle, yaw angle, and roll angle are calculated based on the respective interference coefficients, and the left and right outer blade actuators 13 are independently operated based on the calculated values. By outputting, the attitude control of the aircraft 14 is performed. The attitude of the aircraft is detected by the attitude sensor 15 and fed back to the controller 11 for feedback control. Therefore, the attitude can be controlled by the outer wing 5 and the vertical stabilizer in the conventional aircraft can be eliminated.

  The present invention can eliminate the vertical stabilizer of the aircraft, reduce the weight of the aircraft, and can be applied not only to the delta wing of the supersonic aircraft main wing but also to the delta wing of the subsonic aircraft. Is also applicable.

It is a plane conceptual diagram of the embodiment of the present invention. It is an example of the comparison of the front view of the airframe which raised the conventional delta wing | blade and the outer wing | blade of this invention. It is an example of the operation mechanism concept which changes a mounting angle independently for the left and right outer wings in the present invention. It is an example of the concept of the whole body lift distribution by the variable delta wing | blade in this invention. It is an example of the attitude | position conceptual block diagram by the variable delta wing | blade in this invention.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Delta wing aircraft 2 Airframe 3 Delta wing 4 Inner wing 5 Outer wing 6 Mounting axis 7 Mach line 10 Central processing unit 12 Storage device 13 Actuator 14 Aircraft 15 Attitude control sensor

Claims (4)

  The main wing is based on a triangular plane, and consists of an inner wing and an outer wing, and the outer wing is provided so as to be rotatable with respect to the inner wing about a diagonal line symmetrical to the aircraft axis. Characteristic variable delta wing aircraft.   The variable delta wing aircraft according to claim 1, wherein the mounting angles of the left and right main wing outer wings can be individually varied.   The variable delta wing aircraft according to claim 1, further comprising a fuel tank in the main wing inner wing.   A delta wing as a main wing is composed of an inner wing and an outer wing, and the outer wing is provided so as to be rotatable with respect to the inner wing about an axis that is symmetrical with respect to the axis of the wing. When flying at high subsonic or supersonic speeds, the left and right outer wings are bent upward to prevent air resistance, and the left and right outer wing mounting angles are individually set against changes in the attitude of the aircraft. An aircraft attitude control method for a variable delta wing aircraft, characterized in that the aircraft attitude is controlled by controlling.

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