A method of inhibiting wing rock motion of a flying wing layout aircraft wing

Abstract

The application discloses a method for inhibiting wing roll motion of a flying wing layout aircraft. The method comprises the following steps: determining an angle of attack range of the wing roll motion of the flying wing layout through a free roll motion wind tunnel test; obtaining a position of a reattachment region of a leading edge vortex on an upper surface of the wing through a space flow field display wind tunnel test when the angle of attack of the wing roll motion passes through the space; and determining a gap jet with a proper position and size according to the position of the reattachment region, so that the gap jet generates a suction effect in the roll motion, controls a dynamic breakage and recovery mechanism of the leading edge vortex, and inhibits the wing roll motion of the flying wing layout. The method has the advantages of simple implementation process, obvious inhibition effect and strong engineering practicability.

Description

Technical Field

[0001] This invention relates to a method for suppressing rocking motion of a flying wing aircraft, and more particularly to a method for suppressing rocking motion of a flying wing aircraft using a slit jet, belonging to the field of aerospace technology. Background Technology

[0002] The flying wing design gets its name from its flat, wing-like shape. By eliminating the horizontal and vertical stabilizers, the flying wing design reduces its wetted area, significantly improving aerodynamic and stealth performance. However, this also leads to insufficient longitudinal and directional stability, hindering its early practical application. With the advent of fly-by-wire control systems that resolved the stability issues, the flying wing's superior stealth capabilities have made it highly valued by various countries, as seen in the successful combat deployment of the B-2 stealth bomber. Currently, the flying wing design has become a common configuration for unmanned aerial vehicles (UAVs).

[0003] With the increasing demands for missile maneuverability and battlefield survivability, flying-wing UAVs increasingly require high maneuverability, which is typically achieved through high angle-of-attack post-stall maneuvers. During high angle-of-attack maneuvers, complex vortex separation flows will occur on the leeward side of the flying wing, inducing wing rocking motion, which manifests as uncontrollable limit cycle oscillations around the body axis, severely impacting the high maneuverability and flight safety of the flying wing design. Figure 1 The figure shows the planar shape of a flying wing UAV. Figure 2 and Figure 3 The image shows the large-amplitude wing rocking motion produced by this configuration at a large angle of attack of 26°.

[0004] Extensive research has been conducted both domestically and internationally on the rocking motion of wings with different aerodynamic configurations, such as slender delta wings, rectangular wings, rotor-integrated fuselage-wing-body combinations, and pointed side-edge fuselage-wing-body combinations. The reasons for the rocking motion of wings differ depending on the aerodynamic configuration. For slender delta wings, the leading-edge vortex changes from a symmetrical vortex to an asymmetrical vortex as the angle of attack increases, inducing an initial non-zero roll moment and triggering an initial deviation from zero roll angle motion. The dynamic normal vortex potential hysteresis of the leading-edge asymmetrical vortex sustains the rocking motion. For rotor-integrated fuselage-wing-body combinations, the rotor-integrated fuselage vortex changes from a symmetrical vortex to an asymmetrical vortex as the angle of attack increases, inducing an initial non-zero roll moment in the upper wing surface flow and triggering the rocking motion. The dynamic vortex pattern switching of the asymmetrical fuselage vortex sustains the rocking motion. For the aerodynamic layout of the fuselage-wing-body combination with a pointed side edge, the fuselage vortex with the pointed side edge changes from a symmetrical vortex to an asymmetrical vortex as the angle of attack increases, and an asymmetrical wing vortex is induced on the upper wing surface. The combined action of the asymmetrical fuselage vortex and the asymmetrical wing vortex generates an initial non-zero rolling moment, which triggers rocking motion. The vorticity and vortex potential dynamic hysteresis of the two pairs of asymmetrical vortices together maintain the rocking motion of the wing.

[0005] The different causes of wing rocking motion make it difficult to find universally applicable measures to suppress it. For slender delta wing configurations, rocking motion suppression measures effectively control the normal vortex potential hysteresis of asymmetric leading-edge vortices. For spin-integrated fuselage-wing-body configurations, measures effectively control the dynamic vortex pattern switching mechanism of fuselage asymmetric vortices. For pointed side-edge fuselage-wing-body configurations, measures effectively control the dynamic hysteresis mechanism of vortex quantity and vortex potential of fuselage and wing asymmetric vortices. Existing wing rocking motion suppression measures include rotating nose micro-disturbance particles, nose blowing, and nose leading-edge extensions, which have produced some effect in suppressing wing rocking motion. However, the aerodynamic configuration adapted to the flying wing configuration differs from that of existing suppression measures, and the flow causes of wing rocking motion in flying wing configurations are still under investigation. These measures cannot effectively suppress wing rocking motion in flying wing configurations, necessitating the development of new flow control methods. Summary of the Invention

[0006] The technical problem to be solved by this invention is: in order to solve the problem of rocking motion of flying wing configurations at high angles of attack, a method is proposed to suppress the rocking motion of flying wing configuration aircraft by means of slot jets. Based on controlling the dynamic breaking and recovery mechanism of the leading edge vortex of the flying wing configuration, slot jets are set in appropriate areas of the flying wing configuration. The slot jets generate a suction effect during rocking motion, controlling the dynamic breaking and recovery mechanism of the leading edge vortex, thereby suppressing the rocking motion of the flying wing configuration.

[0007] The above-mentioned objectives of the present invention are mainly achieved through the following technical solutions:

[0008] A method for suppressing rocking motion of a flying wing configuration aircraft involves determining the position of the leading-edge vortex space vortex reattachment region on the upper surface of the wing, and setting a slit within the leading-edge vortex space vortex reattachment region to form a slit jet during rocking motion of the flying wing configuration aircraft.

[0009] In the above-mentioned method for suppressing the rocking motion of a flying wing aircraft, determining the position of the leading-edge vortex reattachment region on the upper surface of the wing includes:

[0010] Establish a free-rocking model of a flying wing layout aircraft;

[0011] A free rocking motion wind tunnel test was conducted on the free rocking motion model of the flying wing configuration aircraft to obtain the roll angle φ of the free rocking motion model with time at different angles of attack, and to determine the angle of attack range in which the wing rocking motion occurs.

[0012] Calculate the equilibrium position and amplitude of the wing rocking motion at different angles of attack within the stated angle of attack range;

[0013] A fixed angle of attack is selected within the angle of attack range. The rocking motion roll angle is obtained based on the equilibrium position and motion amplitude. A space flow field display wind tunnel test is conducted based on the rocking motion roll angle to obtain the leading edge vortex space structure under the fixed angle of attack at different rocking motion roll angles.

[0014] The position of the leading-edge vortex space vortex reattachment region on the upper surface of the wing is determined based on the leading-edge vortex space structure.

[0015] In the above-mentioned method for suppressing the rocking motion of the wings of flying wing aircraft, the established free rocking model of the flying wing aircraft satisfies the following: it can release the roll degree of freedom in wind tunnel testing, the roll axis of the rocking model coincides with the axis of the support rod installed in the wind tunnel test, and under windless and zero angle of attack conditions, the rocking model can roll freely around the roll axis and maintain balance as it is encountered.

[0016] In the above-mentioned method for suppressing the rocking motion of a flying wing aircraft, the free rocking motion wind tunnel test includes: under open wind conditions, adjusting the rocking model to reach a specified angle of attack, releasing the roll degree of freedom of the rocking model, and recording the change curve of the roll angle φ of the rocking model from the release of the roll degree of freedom over time.

[0017] In the above-mentioned method for suppressing wing rocking motion of flying wing aircraft, the angle of attack range in which wing rocking motion occurs is the angle of attack range in which limit cycle oscillation occurs.

[0018] In the above-mentioned method for suppressing wing rocking motion of flying wing aircraft, the calculation of the equilibrium position and motion amplitude of wing rocking motion at different angles of attack within the angle of attack range includes:

[0019] The formula for calculating the equilibrium position is:

[0020] Where, φ eq For the equilibrium position, n is the number of roll angles recorded from the start time to the end time, and φ is the roll angle. i For each roll angle value;

[0021] The formula for calculating the amplitude of motion is:

[0022] Where, φ am For the amplitude of motion, φ max and φ min These are the maximum and minimum roll angles recorded from the start time to the end time, respectively.

[0023] In the above method for suppressing the rocking motion of a flying wing aircraft, the starting time is 5 seconds after the roll degree of freedom is released.

[0024] In the above-mentioned method for suppressing rocking motion of the wings of flying wing aircraft, the rocking motion roll angle is obtained based on the equilibrium position and the motion amplitude, including: selecting a number of angle points as the rocking motion roll angle within a closed interval with the equilibrium position as the center and the motion amplitude as the radius; and / or, densifying the angles near the equilibrium position, that is, reducing the angle interval between the selected angle points.

[0025] The above-mentioned method for suppressing the rocking motion of a flying wing aircraft includes, within a range of 5° to 10° before and after the equilibrium position, an angle interval of 3° to 10°.

[0026] In the above-mentioned method for suppressing the rocking motion of the wings of flying wing aircraft, the angular interval of the angle densification area is 1 to 2°.

[0027] In the above-mentioned method for suppressing the rocking motion of a flying wing aircraft, the length of the gap in the leading edge vortex space vortex reattachment zone is 10% to 100% of the reattachment zone length; and the width is 5% to 100% of the reattachment zone width.

[0028] Compared with the prior art, the present invention has at least the following beneficial effects:

[0029] (1) This invention discloses a method for suppressing rocking motion of a flying wing by using a slot jet. By conducting wind tunnel tests on free rocking motion, the angle of attack range of rocking motion of the flying wing is determined. At the angle of attack of the rocking motion, a spatial flow field display wind tunnel test is conducted to obtain the position of the leading edge vortex reattachment zone on the upper surface of the wing. Based on the position of the reattachment zone, a slot jet with appropriate position and size is determined. The slot jet generates a suction effect during rocking motion, controlling the dynamic rupture and recovery mechanism of the leading edge vortex, thereby suppressing the rocking motion of the flying wing. This method is simple to implement, has a significant suppression effect, and has strong engineering applicability.

[0030] (2) The present invention has found that the main flow that maintains the rocking motion of the flying wing is the dynamic breaking and recovery mechanism of the leading edge vortex of the flying wing. This dynamic breaking and recovery mechanism of the leading edge vortex provides negative damping for the rocking motion of the wing and drives the rocking motion. The present invention is based on controlling the dynamic breaking and recovery mechanism of the leading edge vortex of the flying wing and setting a slot jet in an appropriate area of ​​the flying wing. The slot jet generates a suction effect in the rocking motion, thereby controlling the dynamic breaking and recovery mechanism of the leading edge vortex and suppressing the rocking motion of the flying wing.

[0031] (3) This invention proposes a method for suppressing wing rocking motion in flying wing configurations using slotted jets. When not in the angle-of-attack region where rocking motion occurs, this method is unnecessary and does not affect the normal flight of the flying wing configuration. When wing rocking occurs, opening the slotted jet effectively controls the leading-edge vortex breakup and recovery mechanism, suppressing the wing rocking motion. This invention's suppression method is simple, has a significant effect, and does not bring additional negative impacts, making it highly practical for engineering applications. Attached Figure Description

[0032] Figure 1 This is a schematic diagram of the flying wing layout model and the position of the slit jet in this invention;

[0033] Figure 2 The rocking motion time history curve of a flying wing configuration with a 26° angle of attack during seamless jet flow in an embodiment of the present invention around the negative equilibrium position;

[0034] Figure 3 The rocking motion time history curve of a flying wing configuration with a 26° angle of attack during seamless jet flow in an embodiment of the present invention around the positive equilibrium position;

[0035] Figure 4 This is the time history curve of the rocking motion suppression around the negative equilibrium position for a 26° angle of attack flying wing configuration with a gap jet in an embodiment of the present invention;

[0036] Figure 5 This is the time history curve of the rocking motion suppression around the positive equilibrium position for a flying wing configuration with a 26° angle of attack when there is a gap jet in an embodiment of the present invention.

[0037] 1-Flying wing layout 2-Slot jet Detailed Implementation

[0038] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments:

[0039] This invention, through experiments, reveals that the leading-edge vortex breakup and recovery mechanism drives the rocking motion of flying wing configurations. Near the equilibrium position, the slot jet in the leading-edge vortex reattachment region can effectively control the leading-edge vortex breakup and recovery mechanism, suppressing the rocking motion of flying wing configurations.

[0040] like Figure 1 The diagram shows a flying wing layout model and the location of the slotted jet according to the present invention. A flying wing aircraft 1 has slotted jets 2 installed on its wings. The method of the present invention for suppressing the rocking motion of the flying wing aircraft involves determining the position of the leading-edge vortex reattachment region on the wing surface and setting slots within this region, thereby creating slotted jets during the rocking motion of the flying wing aircraft. The specific steps are as follows:

[0041] Step A: Design a free-form rock model with a flying wing layout.

[0042] In one optional embodiment, the model satisfies the following: it can release the roll degree of freedom in wind tunnel testing; the roll axis of the rocker model coincides with the axis of the support rod installed in the wind tunnel test; and under no wind and zero angle of attack conditions, the rocker model can roll freely around the roll axis and maintain balance as it is encountered.

[0043] Step B: Conduct wind tunnel tests on the free rocking motion model of the flying wing aircraft to obtain the roll angle φ of the free rocking model as a function of time at different angles of attack, and determine the angle of attack range that generates wing rocking motion.

[0044] In one optional embodiment, the free rocking motion wind tunnel test involves adjusting the rocking model to a specified angle of attack under open air conditions, releasing the rolling degree of freedom of the rocking model, and recording the change curve of the rolling angle φ of the rocking model from the release of the rolling degree of freedom over time.

[0045] In one optional embodiment, the angle of attack range that generates wing rocking motion is the angle of attack range in which limit cycle oscillation occurs, and the angle of attack range can be a single continuous range or multiple intermittent continuous ranges.

[0046] Step C: Within the angle of attack range of the wing rocking motion obtained in step B, calculate the equilibrium position and motion amplitude of the wing rocking motion at different angles of attack.

[0047] The calculation of the equilibrium position and amplitude of the wing rocking motion at different angles of attack within the angle of attack range includes:

[0048] The formula for calculating the equilibrium position is:

[0049] Where, φ eq For the equilibrium position, n is the number of roll angles recorded from the start time to the end time, and φ is the roll angle. i For each roll angle value;

[0050] The formula for calculating the amplitude of motion is:

[0051] Where, φ am For the amplitude of motion, φ max and φ min These are the maximum and minimum roll angles recorded from the start time to the end time, respectively.

[0052] In one optional embodiment, the starting time is 5 seconds after the roll degree of freedom is released.

[0053] Step D: Select a fixed angle of attack within the angle of attack range obtained in step B. Obtain the rocking motion roll angle based on the equilibrium position and motion amplitude in step C. Conduct a space flow field display wind tunnel test based on the rocking motion roll angle to obtain the leading edge vortex space structure under a fixed angle of attack at different rocking motion roll angles.

[0054] The process of obtaining the rock and roll motion roll angle based on the equilibrium position and motion amplitude includes: selecting several angle points as the rock and roll motion roll angle within a closed interval centered on the equilibrium position and with the motion amplitude as the radius; and densifying the angles near the equilibrium position, i.e., reducing the angle interval between the selected angle points.

[0055] In one optional embodiment, angle densification is performed within a range of 5° to 10° before and after the equilibrium position; the angle interval range is 3° to 10°. The angle interval of the angle densification area is 1° to 2°.

[0056] Step E: Determine the position of the leading-edge vortex reattachment zone on the upper surface of the wing near the equilibrium position based on the leading-edge vortex space structure in Step D.

[0057] Step F: Based on the location of the reattachment zone on the upper surface of the wing obtained in Step E, determine the location and size of the slot jet. The slot jet is located within the reattachment zone; select appropriate slot width and length to ensure the formation of the slot jet during rocking motion.

[0058] In one optional embodiment, the length of the slit within the reattachment zone of the leading edge vortex space is 10% to 100% of the reattachment zone length, for example, the length is 10% of the reattachment zone length (18 mm); and the width is 5% of the reattachment zone width (1 mm).

[0059] Step G: Set up the slotted jet obtained in step F, penetrating the upper and lower wing surfaces, in the free rocking motion model of the flying wing configuration. Conduct a free rocking motion wind tunnel test in the angle of attack range of the wing rocking motion to obtain the results of suppressing the wing rocking motion of the flying wing configuration aircraft after setting up the slotted jet. This verifies the reliability of the method of the present invention for suppressing the wing rocking motion of flying wing configuration aircraft.

[0060] Unlike the limiting cycle oscillation of a slender delta wing around zero roll angle, the rocking motion of a flying wing is mainly manifested as a limiting cycle oscillation around a positive or negative roll angle, i.e., a rocking motion biased to one side. This invention has found that the dominant flow maintaining the rocking motion of a flying wing may be the dynamic rupture and recovery mechanism of the leading-edge vortex. This mechanism provides negative damping for the rocking motion, driving it. This invention, based on controlling the dynamic rupture and recovery mechanism of the leading-edge vortex in a flying wing, introduces a slotted jet in an appropriate region of the flying wing. The slotted jet generates a suction effect during the rocking motion, controlling the dynamic rupture and recovery mechanism of the leading-edge vortex and suppressing the rocking motion of the flying wing.

[0061] Example

[0062] like Figure 1 The image shows a flying wing aircraft.

[0063] Step 1: Design a free-form rock model with a wing layout, such as... Figure 1 As shown;

[0064] Step 2: Conduct wind tunnel tests on free rock motion to obtain the curves of the model roll angle φ changing with time at different angles of attack, such as... Figure 2 and Figure 3 The curve shown is the change of roll angle over time at an angle of attack of 26°. The angle of attack range that produces wing rocking motion is determined to be 24° to 46°.

[0065] Step 3: In the range of wing rocking motion angles of attack of 24° to 46° obtained in Step 2, calculate the equilibrium position and motion amplitude of the wing rocking motion at different angles of attack. The equilibrium positions of the wing rocking motion at an angle of attack of 26° are -28° and +28°, and the corresponding amplitudes are 24° and 21°.

[0066] Step 4: In the rocking motion angle of attack range of the wing in Step 2, a fixed angle of attack of 26° is selected. Based on the equilibrium position and motion amplitude in Step 3, the rocking motion roll angles φ = -52° to -4° and 7° to 49° are obtained. The spatial flow field display wind tunnel test is conducted to obtain the leading edge vortex space structure at different rocking motion roll angles under a fixed angle of attack of 26°.

[0067] Step 5: Determine the location of the leading-edge vortex reattachment region near the motion equilibrium positions of -28° and +28° in Step 4 on the upper surface of the wing, such as... Figure 1 The location of the slit jet is shown;

[0068] Step 6: Based on the reattachment zone location obtained in Step 5, determine that the position of the slit jet coincides with the reattachment zone and that the slit width is 1mm and the length is 18mm.

[0069] Step 7: Set up the slotted jet obtained in Step 6, penetrating the upper and lower wing surfaces, on the free rocking model of the flying wing configuration. Conduct a free rocking motion wind tunnel test in the angle of attack range of the wing's rocking motion to obtain the results of suppressing the wing's rocking motion after setting up the slotted jet, such as... Figure 4 and Figure 5 As shown.

[0070] The above description is only the best specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the protection scope of the present invention.

[0071] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for suppressing rocking motion of a flying wing aircraft, characterized in that, The position of the leading-edge vortex space vortex reattachment region on the upper surface of the wing of the flying wing configuration aircraft is determined, and a slot is set in the leading-edge vortex space vortex reattachment region so that the flying wing configuration aircraft forms a slot jet during rocking motion. Determining the position of the leading-edge vortex reattachment region on the upper surface of the wing of a flying wing configuration aircraft includes: Establish a free-rocking model of a flying wing layout aircraft; A free-rocking motion wind tunnel test was conducted on the free-rocking model of the flying wing aircraft to obtain the roll angle of the free-rocking model at different angles of attack. The curve of change over time determines the angle of attack range that produces wing rocking motion; Calculate the equilibrium position and amplitude of the wing rocking motion at different angles of attack within the stated angle of attack range; A fixed angle of attack is selected within the angle of attack range. The rocking motion roll angle is obtained based on the equilibrium position and motion amplitude. A space flow field display wind tunnel test is conducted based on the rocking motion roll angle to obtain the leading edge vortex space structure under the fixed angle of attack at different rocking motion roll angles. The position of the leading-edge vortex space vortex reattachment region on the upper surface of the wing is determined based on the leading-edge vortex space structure.

2. The method for suppressing rocking motion of a flying wing aircraft according to claim 1, characterized in that, The established free rocking model of the flying wing aircraft satisfies the following: it can release the roll degree of freedom in wind tunnel testing; the roll axis of the rocking model coincides with the axis of the support rod installed in the wind tunnel test; and under windless and zero angle of attack conditions, the rocking model can roll freely around the roll axis and maintain balance as it goes.

3. The method for suppressing rocking motion of a flying wing aircraft according to claim 1, characterized in that, The free rocking motion wind tunnel test includes: under open wind conditions, adjusting the rocking model to a specified angle of attack, releasing the roll freedom of the rocking model, and recording the roll angle of the rocking model. The curve showing the change over time starting from the release of the roll degree of freedom.

4. The method for suppressing rocking motion of a flying wing aircraft according to claim 1 or 3, characterized in that, The angle of attack range that generates wing rocking motion is the angle of attack range in which the limiting cycle oscillation occurs.

5. The method for suppressing rocking motion of a flying wing aircraft according to claim 1, characterized in that, The calculation of the equilibrium position and motion amplitude of the wing rocking motion at different angles of attack within the angle of attack range includes: The formula for calculating the equilibrium position is: in, For the equilibrium position, This represents the number of roll angles recorded from the start time to the end time. This represents the value for each roll angle; The formula for calculating the amplitude of motion is: in, For the amplitude of motion, and These are the maximum and minimum roll angles recorded from the start time to the end time, respectively.

6. The method for suppressing rocking motion of a flying wing aircraft according to claim 5, characterized in that, The starting time is 5 seconds after the roll degree of freedom is released.

7. The method for suppressing rocking motion of a flying wing aircraft according to claim 1, characterized in that, The rock movement roll angle is obtained based on the equilibrium position and the motion amplitude, including: selecting a number of angle points at certain angular intervals within a closed interval centered on the equilibrium position and with the motion amplitude as the radius, as the rock movement roll angle; and / or, The angle density is increased near the equilibrium position of the motion, that is, the angle interval between the selected angle points is reduced.

8. The method for suppressing rocking motion of a flying wing aircraft according to claim 7, characterized in that, This includes increasing the angle density within a range of 5° to 10° before and after the equilibrium position; The specified angular interval range is 3° to 10°.

9. The method for suppressing rocking motion of a flying wing aircraft according to claim 7 or 8, characterized in that, The angle interval of the angle-encrypted area is 1~2°.

10. The method for suppressing rocking motion of a flying wing aircraft according to claim 1, characterized in that, The length of the gap in the reattachment zone of the leading edge vortex space is 10% to 100% of the length of the reattachment zone; the width is 5% to 100% of the width of the reattachment zone.

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