Tailstock type vertical take-off and landing aircraft with foldable outer-section wings and design method of tailstock type vertical take-off and landing aircraft

Abstract

The invention discloses a tailstock type vertical take-off and landing aircraft with foldable outer-section wings and a design method of the tailstock type vertical take-off and landing aircraft, the outer-section wings can be folded and unfolded by matching a folding and unfolding driving mechanism with a four-bar mechanism, and the weight brought by a folding device is reduced on the premise of ensuring driving force and maintainability; when the wings are completely folded, the aircraft is in a four-rotor state, and when the wings are completely unfolded, the aircraft is in a fixed-wing state; the folding and unfolding of the wings give consideration to the conversion of the flight form, and the folding of the outer-section wings is realized on the premise of ensuring the reasonable gravity center and inertia of the fixed wing state and the rotor wing state at each folding and unfolding angle through the reasonable configuration of the gravity center of each shaft of the vertical take-off and landing aircraft, so that the space cost of the aircraft is reduced.

Description

Technical Field

[0001] This invention relates to the field of foldable wing vertical takeoff and landing (VTOL) aircraft technology, and more particularly to a tail-seat VTOL aircraft with foldable outer wings and its design method. Background Technology

[0002] Mechanical folding and unfolding technology has seen significant development in many fields, especially in aviation and aerospace, achieving efficient folding and unfolding through multi-link hinge systems, hydraulic drives, and lightweight materials. Typical applications include the morphing wingtip of the XB-70 and the foldable wingtip of the Boeing 777X, which employ hinges and hydraulic actuators. The XB-70's morphing wingtip folds downwards to utilize shock waves to increase lift at high speeds, while the Boeing 777X's wingtip unfolds to improve cruise efficiency. However, current technologies still face challenges such as numerous drive units, increased weight, the impact of folding and unfolding on the center of gravity affecting the control of vertical takeoff and landing aircraft during hovering, and the resulting wing shape with angles of attack, dihedral, or anhedral, affecting control performance in fixed-wing cruise states. Summary of the Invention

[0003] Purpose of the invention: This invention provides a tail-seat vertical takeoff and landing aircraft with foldable outer wings and its design method. It achieves folding of the outer wings while ensuring reasonable center of gravity and moment of inertia of the rotor in fixed wing state and at various folding angles, thereby reducing the space cost of the aircraft.

[0004] Technical Solution: The present invention discloses a tail-seat vertical takeoff and landing aircraft with a foldable outer wing, comprising: a fuselage, a vertical tail, and a wing. The wing is divided into an inner wing section and an outer wing section. The inner side of the inner wing section is directly connected to the fuselage. A folding and unfolding drive mechanism is provided on the outer side of the inner wing section to drive the outer wing section to fold and unfold. The folding and unfolding drive mechanism is composed of a drive servo, a servo rocker arm, a carbon linkage, and part of the main spars of the inner and outer wing sections, forming a four-bar linkage. The vertical tail is located at the rear of the fuselage and serves as a landing support during vertical takeoff and landing, and as a vertical tail during level flight.

[0005] Furthermore, the folding and unfolding drive mechanism is driven by a drive servo motor, which drives the servo motor rocker arm and subsequently the carbon linkage, ultimately driving the four-bar linkage with the inner and outer wing main spars to control the folding and unfolding motion of the outer wing.

[0006] Accordingly, a design method for a tail-seat vertical takeoff and landing aircraft with foldable outer wings includes the following steps:

[0007] When fully deployed, the Z-axis coordinate of the outer motor is the same as the Z-axis coordinate of the fuselage center; in other deployment angle states, the Z-axis coordinate of the outer motor is always less than the Z-axis coordinate of the fuselage center; the origin O of the body coordinate system is the intersection of the leading edge of the airfoil of the left and right sides of the aircraft with the plane of symmetry of the aircraft; the X-axis points towards the nose in the plane of symmetry of the aircraft; the Z-axis points towards the belly of the aircraft perpendicular to the X-axis in the plane of symmetry of the aircraft; and the Y-axis points towards the right side of the aircraft perpendicular to the plane of symmetry.

[0008] Using known and estimated geometric data and the weight of key equipment, the Z-axis center of gravity of the whole machine under different folding angles is estimated according to the geometric relationship. The midpoint of the outer wing is used as the reference point. Constraint functions and cost functions are set to find the inner wing dihedral angle and outer wing anhedral angle that minimize the cost function and satisfy the constraint functions.

[0009] By selecting a reasonable dihedral angle for the inner wing section, the Z-axis coordinate of the center of gravity will not deviate too far from the Z-axis coordinate of the intersection point of the diagonal rotor lines during the folding process of the outer wing section. This allows for vertical takeoff from the tailstock under the premise of installing four motors of the same specifications and with the same output. The outer wing section is deployed in rotor mode, and after deployment, the aircraft accelerates by tilting down to transition to level flight. During landing, the aircraft first tilts up to climb and decelerate, and after completing the attitude transition, it descends in the deployed rotor mode. At a low altitude, the outer wing section is folded, and the aircraft lands in this state.

[0010] Furthermore, the following constraint functions are defined:

[0011]

[0012] In the formula, The projected span of a single-segment wing. The Z-axis offset is the position of the battery's center of gravity. The Z-axis offset is the position of the servo motor's center of gravity. The maximum mass of a single-segment wing on one side. For the weight of the fuselage including the equipment, For the single-sided folding drive servo and auxiliary linkage mass, For the mass of a single-sided battery, The weight per unit length of the wing is determined by the process and chord length. First, the actual length of a single wing segment is calculated. The weight of a single wing segment must not exceed the maximum weight limit derived from the overall design. The larger the dihedral angle, the longer and heavier the single wing segment. Then, the Z-axis position of the center of gravity of the important components relative to the midpoint of the outer wing segment is calculated. The closer the center of gravity on the Z-axis is to the midpoint of the outer wing segment, the better.

[0013] Furthermore, based on geometric relationships, the desired anti-reflection angle should be around 30°, and the smaller the better, in terms of the folding angle. The test points were set at 0°, 15°, 30°, 45°, and 60° for different dihedral angles of the inner wing section. The cost function values ​​for different inversion angles were obtained by performing traversal calculations with a step size of 0.1° within the range of 20° to 30°. The cost function is as follows:

[0014]

[0015] in This represents the weighting of the weight of a single wing segment. This represents the weight of a single wing segment. The weights are the values ​​of the selected unfolding angles arranged from smallest to largest. This represents the Z-axis distance between the aircraft's center of gravity and the midpoint of the outer wing section at different folding angles. A lighter single wing section is better, and the closer the aircraft's center of gravity is to the midpoint of the outer wing section at different folding angles, the better. The weights for different folding angles gradually decrease from 0° to 60°. .

[0016] Furthermore, the optimal values ​​are 26.6° for the dihedral angle of the inner wing section and the anhedral angle of the outer wing section.

[0017] Beneficial effects: Compared with the prior art, the present invention has the following significant advantages: the folding and unfolding of the outer wing can be achieved by a folding and unfolding drive device in conjunction with a four-bar linkage, reducing the weight of the folding device while ensuring driving force and maintainability; when the wing is fully folded, the aircraft is in quadrotor mode, and when the wing is fully unfolded, the aircraft is in fixed-wing mode; the wing folding and unfolding takes into account the change of flight mode, and by rationally configuring the center of gravity of each axis of the vertical take-off and landing aircraft, the folding of the outer wing is achieved while ensuring reasonable center of gravity and moment of inertia in fixed-wing mode and rotor mode at various folding angles, thereby reducing the space cost of the aircraft. Attached Figure Description

[0018] Figure 1 This is a schematic diagram of the fully deployed outer wing structure of the present invention.

[0019] Figure 2 This is a schematic diagram of the outer wing structure of the present invention when fully retracted.

[0020] Figure 3 This is a diagram showing the state of the folding mechanism of the present invention when the outer section of the wing is fully deployed.

[0021] Figure 4 This is a diagram showing the state of the folding mechanism of the present invention when the outer section of the wing is fully retracted.

[0022] Figure 5 This is a simplified geometric diagram illustrating the relationship between the inner and outer wing sections and the fuselage according to the present invention.

[0023] Figure 6This is a graph showing the cost function variation relationship under different dihedral angles of the inner wing section and the anhedral angle of the outer wing section according to the present invention.

[0024] Figure 7 This is a graph showing the variation of the aileron control derivative under different flapping angles according to the present invention.

[0025] Among them, 101. Fuselage; 102. Right inner wing section; 103. Right outer wing section; 104. Vertical tail; 105. Left inner wing section; 106. Left outer wing section; 107. Right elevon; 108. Left elevon; 109. Motor mount 1; 110. Motor mount 2; 111. Motor mount 3; 112. Motor mount 4; 113. Cover plate; 114. Wing rib; 115. Main beam connection point; 11 6. Front wall connection point; 117. Rear wall connection point; 118. Main beam pivot; 119. Front wall pivot; 120. Rear wall pivot; 121. Drive servo; 122. Servo arm; 123. Carbon fiber tie rod; 124. Right inner wing main spars; 125. Right outer wing main spars; 126. Right inner wing front wall; 127. Right outer wing front wall; 128. Right inner wing rear wall; 129. Right outer wing rear wall. Detailed Implementation

[0026] like Figures 1-4 As shown, a tail-seat vertical takeoff and landing aircraft with a foldable outer wing includes a fuselage, a vertical tail, and a wing. The wing is divided into an inner wing section and an outer wing section. The inner side of the inner wing section is directly connected to the fuselage. A folding and unfolding drive mechanism is provided on the outer side of the inner wing section to drive the outer wing section to fold and unfold. The folding and unfolding drive mechanism consists of a drive servo, a servo rocker arm, a carbon linkage, and part of the main spars of the inner and outer wing sections, forming a four-bar linkage. The vertical tail is located at the rear of the fuselage and serves as a landing support during vertical takeoff and landing, and as a vertical tail during level flight.

[0027] The front walls of the right inner wing section 102 and the right outer wing section 103 are connected by a front wall pivot 119. The front walls 126 and 127 of the right inner wing section and the right outer wing section are connected by a main wall pivot 118. The rear walls 128 and 129 of the right inner wing section and the right outer wing section are connected by a rear wall pivot 120. The front walls of the left and right inner wings are fixedly connected by a front wall connection point 116. The main beams of the left and right inner wings are fixedly connected by a mortise and tenon joint at a main beam connection point 115. The rear walls of the left and right inner wings are fixedly connected by a connection point 117. After the front wall, main spars, and rear wall of each inner wing section on both sides are fixed together as a single unit at connection points 116, 115, and 117 inside the fuselage, they are interconnected by mortise and tenon joints via wing ribs 114, thus connecting the front wall, main spars, and rear wall of the inner wing section. A cover plate 113 further connects the front wall, main spars, and wing ribs of each inner wing section, forming a complete inner wing section. The outer wing structure is similar. Taking the right outer wing section as an example, the front wall 127, main spars 125, and rear wall 129 of the outer wing section are interconnected by mortise and tenon joints via wing ribs, and the front wall, main spars, and wing ribs of the outer wing section are fixed together via a cover plate. The main spars connection point 115, the front wall connection point 116, and the rear wall connection point 117 traverse the fuselage 101 and are fixed to the fuselage.

[0028] The folding and unfolding drive mechanism is driven by the drive servo motor 121, which drives the servo motor rocker arm 122 and subsequently the carbon linkage 123, ultimately driving the four-bar linkage with the inner wing main spars 124 and the outer wing main spars 125, thus controlling the folding and unfolding motion of the outer wing.

[0029] The front and rear walls, together with the main beam, form a rotation axis that runs longitudinally through the entire wing chord, including the front wall rotation axis 119, the main beam rotation axis 118, and the rear wall rotation axis 120. The outer section of the wing rotates around this rotation axis.

[0030] 104 is the vertical stabilizer fixed to the tail of the fuselage; 105 is the left inner wing section; 106 is the left outer wing section; 107 is the right elevon; 108 is the left elevon. All elevons are fixed to the rear wall of the outer wing section and are used to generate control torque. 109 is the mounting bracket for motor 1; 110 is the mounting bracket for motor 2; 111 is the mounting bracket for motor 3; and 112 is the mounting bracket for motor 4. All motor mounting brackets are fixed to the main spars and front wall of the outer wing section.

[0031] In the fully deployed state, the outermost motor has the same Z-axis coordinate as the fuselage. At other deployment angles, the Z-axis coordinate of the outermost motor is always less than the Z-axis coordinate of the fuselage center. The origin O of the fuselage coordinate system is the intersection of the leading edges of the wing airfoils on both sides of the aircraft and the aircraft's plane of symmetry. The X-axis points towards the nose within the plane of symmetry, the Z-axis points towards the belly of the aircraft perpendicular to the X-axis within the plane of symmetry, and the Y-axis points towards the right side of the aircraft perpendicular to the plane of symmetry.

[0032] After the preliminary design was completed, using the known and estimated geometric data shown in Table 1, the weight of key equipment, and their geometric relationships, the Z-axis center of gravity of the entire aircraft at different folding angles was estimated. Using the midpoint of the outer wing section as a reference point, constraint functions and cost functions were set, and the anhedral angle of the inner wing section and the anhedral angle of the outer wing section that minimized the cost function and satisfied the constraint functions were sought. Since the motor, ESC, propeller, and servo surfaces are all located on the outer wing section and are symmetrical about the midpoint of the outer wing section, they are not considered in the discussion of equipment distribution. A simplified schematic diagram of the wing folding is shown below. Figure 5 As shown, the following constraints are set:

[0033]

[0034] In the formula, The projected span of a single-segment wing. The Z-axis offset is the position of the battery's center of gravity. The Z-axis offset is the position of the servo motor's center of gravity. The maximum mass of a single-segment wing on one side. For the weight of the fuselage including the equipment, For the single-sided folding drive servo and auxiliary linkage mass, For the mass of a single-sided battery, This refers to the weight per unit length of the wing under this process and chord length. The constraints are as follows: First, calculate the actual length of a single wing segment. The calculated weight of a single wing segment must not exceed the maximum weight limit derived from the overall design. The larger the dihedral angle, the longer the single wing segment, and the greater the weight. Then, calculate the Z-axis position of the center of gravity of the important components relative to the midpoint of the outer wing segment. The closer the distance between the center of gravity on the Z-axis and the midpoint of the outer wing segment, the better.

[0035] Table 1. Known and Estimated Geometric Data

[0036]

[0037] Based on simple geometric relationships, the desired anti-reflection angle should be around 30°, and the smaller the better, based on the folding angle. The test points were set at 0°, 15°, 30°, 45°, and 60° for different dihedral angles of the inner wing section. The cost function values ​​for different inversion angles were obtained by performing traversal calculations with a step size of 0.1° within the range of 20° to 30°. The cost function is as follows:

[0038]

[0039] in This represents the weighting of the weight of a single wing segment. This represents the weight of a single wing segment. The weights are the values ​​of the selected unfolding angles arranged from smallest to largest. This represents the Z-axis distance between the aircraft's center of gravity and the midpoint of the outer wing section at different folding angles. A lighter single wing section is better, and the closer the aircraft's center of gravity is to the midpoint of the outer wing section at different folding angles, the better, but their importance varies. Since different dihedral angles have a limited impact on the length increase of a single wing section, their effect on weight is relatively small, while their impact on the center of gravity is significant. Furthermore, since the primary focus is on the center of gravity when the outer wing section is fully retracted, the weights for different folding angles gradually decrease from 0° to 60°. The change in the cost function value is as follows: Figure 6 As shown, the optimal values ​​are 26.6° for the dihedral angle of the inner wing section and the anhedral angle of the outer wing section.

[0040] Figure 1 The state shown is when fully deployed. When the deployment angle is 63.4°, the four motors are arranged in an isosceles trapezoidal shape on the YOZ plane. Figure 2 When fully folded and with a folding angle of 0°, the four motors are arranged in a rectangular pattern on the YOZ plane.

[0041] The outer end of the fixed wing sparb is fixed with a drive mechanism, which is driven by a servo motor. This mechanism drives a four-bar linkage consisting of a servo motor rocker arm, a carbon tie rod, and a portion of the main spar of the inner and outer wing sections, controlling the folding and unfolding of the outer wing section.

[0042] The folding drive mechanism is attached to the main spar of the inner wing on the outer side of the wing. The pull rod and rocker arm are very small, which facilitates maintenance, ensures control and strength, and does not increase drag too much.

[0043] By selecting a reasonable dihedral angle for the inner wing section, the Z-axis coordinate of the center of gravity is ensured not to deviate too far from the Z-axis coordinate of the intersection point of the diagonal rotor lines during the folding process of the outer wing section. The relationship between the center of gravity and the motor center is shown in Table 2. This allows for vertical takeoff from the tailstock under the premise of installing four motors of the same specification and with the same output. The outer wing section is deployed in rotor mode, and after deployment, the aircraft accelerates down to level flight. During landing, the aircraft first climbs and decelerates to complete the attitude transition. After completing the attitude transition, it descends in the deployed rotor mode. At a low altitude, the outer wing section is folded, and the aircraft lands in this state.

[0044] Table 2 Relationship between center of gravity and motor center

[0045]

[0046] like Figure 7 As shown, throughout the entire folding and unfolding process, under normal operating airflow angles, this tail-seat VTOL aircraft can consistently maintain roll self-stability in fixed-wing coordinates. Roll self-stability refers to the aircraft's ability to generate a restoring moment after being subjected to roll disturbances, causing it to tend towards its original equilibrium state. This stability is mainly determined by the roll moment coefficient. Side slip angle derivative Decide, This represents the aircraft's roll self-stabilization. It can also adapt to another transition mode: deploying while nose-down and decelerating while nose-up and retracting. Maintaining roll self-stabilization during these processes greatly reduces control difficulty.

Claims

1. A tail-seat vertical takeoff and landing aircraft with foldable outer wings, characterized in that, include: The fuselage, vertical tail, and wings are divided into an inner wing section and an outer wing section. The inner wing section is directly connected to the fuselage on the inside. The outer side of the inner wing section is equipped with a folding and unfolding drive mechanism to drive the outer wing section to fold and unfold. The folding and unfolding drive mechanism consists of a drive servo, a servo rocker arm, a carbon linkage, and part of the main spars of the inner and outer wing sections, forming a four-bar linkage. The vertical tail is located at the rear of the fuselage and serves as a landing support during vertical takeoff and landing, and as a vertical tail fin during level flight.

2. The tail-seat vertical takeoff and landing aircraft with foldable outer wings as described in claim 1, characterized in that, The folding and unfolding drive mechanism is driven by a drive servo motor, which drives the servo motor rocker arm and then the carbon linkage, ultimately driving the four-bar linkage with the inner and outer wing main spars to control the folding and unfolding motion of the outer wing.

3. A design method for a tail-seat vertical takeoff and landing aircraft with foldable outer wings as described in claim 1, characterized in that, Includes the following steps: When fully deployed, the Z-axis coordinate of the outer motor is the same as the Z-axis coordinate of the fuselage center; in other deployment angle states, the Z-axis coordinate of the outer motor is always less than the Z-axis coordinate of the fuselage center; the origin O of the body coordinate system is the intersection of the leading edge of the airfoil of the left and right sides of the aircraft with the plane of symmetry of the aircraft; the X-axis points towards the nose in the plane of symmetry of the aircraft; the Z-axis points towards the belly of the aircraft perpendicular to the X-axis in the plane of symmetry of the aircraft; and the Y-axis points towards the right side of the aircraft perpendicular to the plane of symmetry. Using known and estimated geometric data and the weight of key equipment, the Z-axis center of gravity of the whole machine under different folding angles is estimated according to the geometric relationship. The midpoint of the outer wing is used as the reference point. Constraint functions and cost functions are set to find the inner wing dihedral angle and outer wing anhedral angle that minimize the cost function and satisfy the constraint functions. By selecting a reasonable dihedral angle for the inner wing section, the Z-axis coordinate of the center of gravity will not deviate too far from the Z-axis coordinate of the intersection point of the diagonal rotor lines during the folding process of the outer wing section. This allows for vertical takeoff from the tailstock under the premise of installing four motors of the same specifications and with the same output. The outer wing section is deployed in rotor mode, and after deployment, the aircraft accelerates by tilting down to transition to level flight. During landing, the aircraft first tilts up to climb and decelerate, and after completing the attitude transition, it descends in the deployed rotor mode. At a low altitude, the outer wing section is folded, and the aircraft lands in this state.

4. The design method of the tail-seat vertical takeoff and landing aircraft with foldable outer wings as described in claim 3, characterized in that, Define the following constraint functions: In the formula, The projected span of a single-segment wing. The Z-axis offset is the position of the battery's center of gravity. The Z-axis offset is the position of the servo motor's center of gravity. The maximum mass of a single-segment wing on one side. For the weight of the fuselage including the equipment, For the single-sided folding drive servo and auxiliary linkage mass, For the mass of a single-sided battery, The weight per unit length of the wing is determined by the process and chord length. First, the actual length of a single wing segment is calculated. The weight of a single wing segment must not exceed the maximum weight limit derived from the overall design. The larger the dihedral angle, the longer and heavier the single wing segment. Then, the Z-axis position of the center of gravity of the important components relative to the midpoint of the outer wing segment is calculated. The closer the center of gravity on the Z-axis is to the midpoint of the outer wing segment, the better.

5. The design method of the tail-seat vertical takeoff and landing aircraft with foldable outer wings as described in claim 3, characterized in that, Based on geometric relationships, the desired anti-reflection angle should be around 30°, and the smaller the better, in terms of the folding angle. The test points were set at 0°, 15°, 30°, 45°, and 60° for different dihedral angles of the inner wing section. The cost function values ​​for different anti-reverse angles are obtained by performing traversal calculations with a step size of 0.1° within the range of 20° to 30°.

6. The design method of the tail-seat vertical takeoff and landing aircraft with foldable outer wings as described in claim 5, characterized in that, The cost function is as follows: in This represents the weighting of the weight of a single wing segment. This represents the weight of a single wing segment. The weights are the values ​​of the selected unfolding angles arranged from smallest to largest. This represents the Z-axis distance between the aircraft's center of gravity and the midpoint of the outer wing section at different folding angles. A lighter single wing section is better, and the closer the aircraft's center of gravity is to the midpoint of the outer wing section at different folding angles, the better. The weights for different folding angles gradually decrease from 0° to 60°. .

7. The design method of the tail-seat vertical takeoff and landing aircraft with foldable outer wings as described in claim 6, characterized in that, The optimal angles are 26.6° for the dihedral angle of the inner wing section and the anhedral angle of the outer wing section.

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