A jet flow control device for aircraft model flight test and a test method thereof

Abstract

The invention discloses a jet flow control device for aircraft model flight test and a test method thereof, and belongs to the field of aircraft test. The control device includes a driving motor, a motor mounting seat, a coupling, a valve adapter shaft, a valve mounting seat, a valve body and a valve core, and the valve core is a V-mouth ball core whose inlet area does not change with the rotation of the ball core. Based on this device, a use strategy of jet control and traditional control rudder combination is developed. Compared with the traditional proportional flow control device, while ensuring the high-frequency and large-flow proportional control requirements, the energy loss is lower, the proportional control response speed is higher, and the pipeline flow stability is better. The strict requirements of load weight volume and jet control accuracy during aircraft model flight test can be met, and the control efficiency of jet control measures can be verified by aircraft model flight test.

Description

Technical Field

[0001] The invention belongs to the field of aircraft testing, and in particular relates to an aircraft model flight test jet flow control device and a test method thereof. Background Art

[0002] As the application scenarios of aircraft change, higher requirements are placed on the performance of aircraft. The control capability of traditional aircraft rudders can no longer meet the requirements of aircraft takeoff, green flight and stealth. In this context, jet control measures have become one of the advanced aerodynamic control measures that supplement and enhance, or even replace traditional rudders. As a new type of aerodynamic control measure, the mature application of jet control measures needs to be verified through a large number of aircraft model atmospheric and wind tunnel flight tests. Under the conditions that the model meets the dynamics and size similarity, the flow control device of the real aircraft jet control measures cannot be applied to the model flight test due to weight and volume restrictions. The existing small flow control devices also have problems such as large volume, low response frequency, small flow, poor linearity of proportional control, and large momentum of post-valve flow artery, which cannot meet the requirements of aircraft model flight test verification. Therefore, it is necessary to develop a jet flow control device for aircraft model flight test to improve the flight test verification capability of aircraft model. Summary of the invention

[0003] In view of the above shortcomings, the present invention provides a jet flow control device for aircraft model flight tests, which can solve the problems of large volume and weight of the flow control device while meeting the conditions of large flow, high frequency response, low pulsation and proportional control performance, and realize the closed-loop control of the flight attitude of the aircraft model using jets.

[0004] The technical scheme adopted by the present invention is as follows: a jet flow control device for an aircraft model flight test, comprising a drive motor, a motor mounting seat, a coupling, a valve adapter shaft, a valve mounting seat, a valve body and a valve core; the drive motor is mounted on the motor mounting seat, the output shaft of the drive motor is connected to the valve adapter shaft through a coupling, the valve body is connected to the valve mounting seat, the valve adapter shaft is connected to the valve body, the valve core is located in the valve body, the two ends of the valve body are respectively a direct current flow channel inlet and an outlet, the valve core is a hollow sphere, and one side of the valve core is provided with a cross section for a horizontal The triangular cross-section outlet is connected to the outlet of the DC flow channel, and the surface of the other side of the valve core is opened with a horizontal, strip-shaped long opening with an arc of 90° in the counterclockwise direction with the center of the valve core as the center and is connected to the inlet of the DC flow channel. The width of the strip-shaped long opening is equal to the inner diameter of the DC flow channel. When the triangular cross-section outlet gradually enters the DC flow channel, the cross-sectional area of ​​the triangular cross-section outlet in contact with the DC flow channel gradually increases, and during the process of the valve core rotating and opening in the valve body, the cross-sectional shape of the strip-shaped long opening in contact with the DC flow channel does not change with the rotation of the valve core.

[0005] Furthermore, the rotation angle range of the valve core in the valve body is 0°-90°

[0006] Furthermore, the triangular cross-section outlet is an isosceles triangle, and the vertex angle thereof is 30°-60°.

[0007] Furthermore, the driving motor is a brushless motor, whose stall torque is greater than the valve core starting torque, the rotation angular velocity is not less than 300 degrees / second, and the signal frequency is 50Hz.

[0008] Another object of the present invention is to provide a test method for a jet flow control device for an aircraft model flight test, which is implemented based on the above-mentioned jet flow control device for an aircraft model flight test and includes the following steps:

[0009] Step 1: Install the flow control device inside the aircraft model, connect the inlet of the flow control device to the air supply pipeline, and connect the outlet of the flow control device to the ejector;

[0010] Step 2: Measure the momentum coefficient C of the valve of the flow control device at different openings μ , the formula is as follows:

[0011]

[0012] In the formula, q ∞ is the dynamic pressure of the aircraft model during flight; S is the reference area of ​​the model; the gas mass flow rate in the pipeline Ejector outlet velocity u j

[0013] Step 3: Conduct a static force test on the model in a wind tunnel to measure the effect of the jet on the six-element aerodynamic coefficient under different momentum coefficients;

[0014] Step 4: According to the influence of momentum coefficient on aerodynamic coefficient under different valve openings, establish the rudder trim distribution strategy of coupled jet control. The specific trim formula is as follows:

[0015] Outer aileron allocation formula:

[0016]

[0017] Inner aileron allocation formula:

[0018] δ e,right =δ e,left =δ q (3)

[0019] Rudder and jet combination, distribution formula:

[0020]

[0021] Where, k1 is the rudder surface allocation ratio; k2 is the rudder surface-jet conversion ratio; δ a,left ,δ a,right They are the left and right outer ailerons respectively; δ e,left ,δ e,right They are the left and right inner ailerons respectively; δ wt,left ,δ wt,right They are the left and right wing rudders respectively; M sl,left 、M sl,right are the left and right jet momentum respectively; δ p ,δ q ,δ r They are the actual control amount of roll, the actual control amount of pitch, and the actual control amount of yaw;

[0022] Step 5: Design the control law based on the control surface allocation strategy. The control law architecture adopts the structure of inner and outer loops. The outer loop control is position loop control, which gives pitch angular velocity, yaw angular velocity and roll angle targets for angle of attack, sideslip angle and roll angle targets. The inner loop control is speed loop control, which gives deflection angle instructions for inner and outer ailerons and rudder for angular velocity targets. Both the inner and outer loops adopt feedback control methods.

[0023] The outer loop control first obtains the angle of attack velocity target, sideslip angle velocity target and roll angle velocity target according to the angle of attack, sideslip angle and roll angle errors. The formula is as follows:

[0024]

[0025] Among them, α, α c are the angle of attack and the angle of attack target, respectively; β, β c are the sideslip angle and the sideslip angle target respectively; φ, φ c are roll angle and roll angle target respectively; are the angle of attack velocity target, sideslip angular velocity target, and roll angular velocity target, respectively. α , K β , K φ are the angle of attack control gain, sideslip angle control gain, and roll angle control gain respectively;

[0026] The sideslip angular velocity is decomposed into the roll axis and yaw axis through the angle of attack, and the formula for obtaining the body axis angular velocity target is as follows:

[0027]

[0028] Among them, p c ,q c 、r c They are the roll axis angular velocity target, pitch axis angular velocity target, and yaw axis angular velocity target of the inner loop, which are used for inner loop control;

[0029] The inner loop control performs rudder control according to the angular velocity target. The formula for the rudder control increment is as follows:

[0030]

[0031] Among them, p, q, and r are the angular velocity of the roll axis, the angular velocity of the pitch axis, and the angular velocity of the yaw axis, respectively. p , Δδ q , Δδ r are the roll control increment, pitch control increment, and yaw control increment, respectively, K p , K q , K r They are roll control gain, pitch control gain, and yaw control gain, respectively.

[0032] The formula for adding the control increment and the control trim amount to obtain the actual control amount is as follows:

[0033]

[0034] Among them, δ p,0 , δ q,0 , δ r,0 They are roll trim, pitch trim, and yaw trim respectively;

[0035] Step 6: Based on the control surface allocation strategy of step 4 and the control law of step 5, conduct a model flight test to verify the flight control capability of the ejector.

[0036] Advantages and beneficial effects of the present invention: Compared with traditional proportional flow control devices, the present invention can ensure high-frequency and large-flow proportional control requirements while having lower energy loss, higher proportional control response speed, and better pipeline flow stability. It can meet the stringent requirements of load weight volume and jet control accuracy during aircraft model flight tests, and can be widely used in aircraft model flight tests to enhance the ability of aircraft model flight tests to verify the control effectiveness of jet control measures. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Figure 1 It is a structural diagram of the jet flow control device of the present invention;

[0038] Figure 2 It is a structural diagram of the valve core of the jet flow control device of the present invention;

[0039] Figure 3 It is a cross-sectional view of the valve core of the jet flow control device of the present invention;

[0040] Figure 4 This is a flow control effect comparison curve between the present invention and the traditional proportional valve;

[0041] Figure 5This is a cloud diagram of the pressure field of the flow simulation in the flow channel of a traditional proportional valve;

[0042] Figure 6 It is a cloud diagram of the pressure field of flow simulation in the flow channel of the present invention;

[0043] Figure 7 This is the cloud diagram of the flow velocity field in the flow channel of a traditional proportional valve;

[0044] Figure 8 It is a cloud diagram of the flow velocity field in the flow channel of the present invention;

[0045] Fig. 9 A curve diagram showing the change in opening of the jet flow control device when the jet flow control device of the present invention is used to carry out a jet control wind tunnel flight verification test;

[0046] Fig.10 A curve diagram showing the change of rudder surface deflection when the jet flow control device of the present invention is used to carry out a jet control wind tunnel flight verification test;

[0047] Fig.11 The present invention is a graph showing the change in heading attitude angle when the jet flow control device of the present invention is used to carry out a jet control wind tunnel flight verification test. DETAILED DESCRIPTION

[0048] In order to make the purpose, technical scheme and advantages of the present invention clearer, the present invention is described below by the specific embodiments shown in the accompanying drawings. However, it should be understood that these descriptions are only exemplary and are not intended to limit the scope of the present invention. In addition, in the following description, the description of well-known structures and technologies is omitted to avoid unnecessary confusion of the concept of the present invention.

[0049] Example 1

[0050] like Figure 1-3As shown, a jet flow control device for an aircraft model flight test comprises a driving motor 1, a motor mounting seat 2, a coupling 3, a valve adapter shaft 4, a valve mounting seat 5, a valve body 6 and a valve core 7; the driving motor 1 is mounted on the motor mounting seat 2, the output shaft of the driving motor 1 is connected to the valve adapter shaft 4 through the coupling 3, the valve body 6 is connected to the valve mounting seat 5, the valve adapter shaft 4 is connected to the valve body 6, the valve core 7 is located in the valve body 6, the two ends of the valve body 6 are respectively a direct current flow channel inlet and outlet, the valve core 7 is a hollow sphere, and an outlet 8 with a cross-section of a horizontally placed isosceles triangle is opened on one side of the valve core 7 to communicate with the direct current flow channel outlet, the isosceles triangle The vertex angle 9 of the cross-sectional outlet 8 is 30°. On the other side of the valve core 7, a horizontal strip-shaped long opening 10 with an arc of 90° is opened on the surface of the valve core in the counterclockwise direction with the ball center of the valve core as the center and is connected to the inlet of the direct current flow channel. The width of the strip-shaped long opening 10 is equal to the inner diameter of the direct current flow channel. The rotation angle range of the valve core 7 in the valve body is 0°-90°. When the triangular cross-sectional outlet 8 gradually enters the direct current flow channel, the cross-sectional area of ​​the isosceles triangular cross-sectional outlet 8 in contact with the direct current flow channel gradually increases, and during the process of the valve core 7 rotating and opening in the valve body, the cross-sectional shape of the strip-shaped long opening 10 in contact with the direct current flow channel does not change with the rotation of the valve core 7. The drive motor 1 adopts a brushless motor, the stall torque is greater than the starting torque of the V-mouth ball valve, the angular motion range is: 0°-90°, the angular velocity is 460 degrees / second, the drive signal frequency is 50Hz, and the size of the electric drive device is 45mm×42mm×21mm. The diameter of the semicircular arc of the left end 11 and the right end 12 of the long strip opening is the same as the inner diameter of the direct current flow channel 13. During the rotation and opening process of the valve core 7 in the valve body, the vertex 9 of the triangular cross-section outlet 8 gradually enters the direct current flow channel, and the cross-sectional area of ​​the direct current flow channel outlet increases. When the triangular cross-section outlet 8 completely enters the direct current flow channel, the flow channel section between the valve core 7 and the valve body 6 is an isosceles triangle; during the rotation and opening process of the valve core 7 in the valve body 6, the two ends of the strip long opening 10 will never enter the direct current flow channel, and the flow channel section between the valve core 7 and the valve body 6 is always a circle with the same diameter as the inner wall of the pipeline, ensuring that the inlet cross-sectional shape does not change with the rotation of the valve core 7, and the inner diameter of the direct current flow channel is 15mm. Two flow control devices are installed inside the aircraft model, the inlet of the flow control device is connected to the air supply pipeline, and the outlets of the two flow control devices are respectively connected to the ejectors on the left and right sides of the model.

[0051] like Figure 4 As shown, the flow rate proportional control effect of the present invention is better than that of the traditional proportional valve, and the higher linearity of the present invention is conducive to the development of the ejector control law. Figure 5 , Figure 6 , Figure 7 , Figure 8As shown, the inlet cross-sectional area of ​​the control valve of the present invention is equivalent to the cross-sectional area of ​​the incoming flow pipeline, there is no process of fluid expansion and contraction, energy loss is reduced, the energy consumption of the air supply system is reduced, the flow velocity in the pipeline is faster, the flow control response speed is improved, the pressure and velocity field pulsation of the flow in the pipeline is reduced, the stability of the flow in the pipeline is increased, and the control accuracy of the jet controller is improved.

[0052] Example 2

[0053] A test method for a jet flow control device for an aircraft model flight test is implemented using the aircraft model flight test jet flow control device described in Example 1, and includes the following steps:

[0054] Step 1: Install the flow control device inside the aircraft model, connect the inlet of the flow control device to the air supply pipeline, and connect the outlet of the flow control device to the ejector;

[0055] Step 2: Measure the momentum coefficient C of the valve of the flow control device at different openings μ , the formula is as follows:

[0056]

[0057] In the formula, q ∞ is the dynamic pressure of the aircraft model during flight; S is the reference area of ​​the model; the gas mass flow rate in the pipeline Ejector outlet velocity u j

[0058] Step 3: Conduct a static force test on the model in a wind tunnel to measure the effect of the jet on the six-element aerodynamic coefficient under different momentum coefficients;

[0059] Step 4: According to the influence of momentum coefficient on aerodynamic coefficient under different valve openings, establish the rudder trim distribution strategy of coupled jet control. The specific trim formula is as follows:

[0060] Outer aileron allocation formula:

[0061]

[0062] Inner aileron allocation formula:

[0063] δ e,right =δ e,left =δ q (3)

[0064] Rudder and jet combination, distribution formula:

[0065]

[0066] Where, k1 is the rudder surface allocation ratio; k2 is the rudder surface-jet conversion ratio; δ a,left , δ a,right They are the left and right outer ailerons respectively; δ e,left , δ e,right They are the left and right inner ailerons respectively; δ wt,left , δ wt,right They are the left and right wing rudders respectively; M sl,left 、M sl,right are the left and right jet momentum respectively; δ p , δ q , δ r They are the actual control amount of roll, the actual control amount of pitch, and the actual control amount of yaw;

[0067] Step 5: Design the control law based on the control surface allocation strategy. The control law architecture adopts the structure of inner and outer loops. The outer loop control is position loop control, which gives pitch angular velocity, yaw angular velocity and roll angle targets for angle of attack, sideslip angle and roll angle targets. The inner loop control is speed loop control, which gives deflection angle instructions for inner and outer ailerons and rudder for angular velocity targets. Both the inner and outer loops adopt feedback control methods.

[0068] The outer loop control first obtains the angle of attack velocity target, sideslip angle velocity target and roll angle velocity target according to the angle of attack, sideslip angle and roll angle errors. The formula is as follows:

[0069]

[0070] Among them, α, α c are the angle of attack and the angle of attack target, respectively; β, β c are the sideslip angle and the sideslip angle target respectively; φ, φ c are roll angle and roll angle target respectively; are the angle of attack velocity target, sideslip angular velocity target, and roll angular velocity target, respectively. α , K β , K φ are the angle of attack control gain, sideslip angle control gain, and roll angle control gain respectively;

[0071] The sideslip angular velocity is decomposed into the roll axis and yaw axis through the angle of attack, and the formula for obtaining the body axis angular velocity target is as follows:

[0072]

[0073] Among them, p c ,q c 、r c They are the roll axis angular velocity target, pitch axis angular velocity target, and yaw axis angular velocity target of the inner loop, which are used for inner loop control;

[0074] The inner loop control performs rudder control according to the angular velocity target. The formula for the rudder control increment is as follows:

[0075]

[0076] Among them, p, q, and r are the angular velocity of the roll axis, the angular velocity of the pitch axis, and the angular velocity of the yaw axis, respectively. p , Δδ q , Δδ r are the roll control increment, pitch control increment, and yaw control increment, respectively, K p , K q , K r They are roll control gain, pitch control gain, and yaw control gain, respectively.

[0077] The formula for adding the control increment and the control trim amount to obtain the actual control amount is as follows:

[0078]

[0079] Among them, δ p,0 , δ q,0 , δ r,0 They are roll trim, pitch trim, and yaw trim respectively;

[0080] Step 6: Based on the control surface allocation strategy of step 4 and the control law of step 5, conduct a model flight test to verify the flight control capability of the ejector.

[0081] like Fig. 9 , Fig.10 , Fig.11 As shown in the figure, when there is no jet control, the heading attitude angle and the deflection of the rudder surface of the aircraft model are in an oscillating state with a large oscillation amplitude. When it is controlled by pure jet, the oscillation amplitude of the heading attitude angle and the deflection of the rudder surface of the aircraft model is significantly reduced to a negligible state. It can be seen that the use of the jet flow control device and the test method can meet the requirements of the aircraft scale model for the volume, weight, flow rate and frequency response of the jet control device, realize the virtual flight test of pure jet heading stabilization control without rudder surfaces, and verify the flight control effect of the ejector.

Claims

1. A test method for a jet flow control device for an aircraft model flight test, which is implemented by using a jet flow control device for an aircraft model flight test, the device comprising a drive motor (1), a motor mounting seat (2), a coupling (3), a valve adapter shaft (4), a valve mounting seat (5), a valve body (6) and a valve core (7); the drive motor (1) is mounted on the motor mounting seat (2), the output shaft of the drive motor (1) is connected to the valve adapter shaft (4) through the coupling (3), the valve body (6) is connected to the valve mounting seat (5), the valve adapter shaft (4) is connected to the valve body (6), the valve core (7) is located in the valve body (6), the two ends of the valve body (6) are respectively a direct current flow channel inlet and an outlet, and the valve core (7) ) is a hollow sphere, a triangular cross-section outlet (8) is opened on one side surface of the valve core (7) and is connected to the outlet of the direct current flow channel, and a horizontal strip-shaped long opening (10) with an arc of 90° is opened on the surface of the other side of the valve core (7) along the counterclockwise direction with the center of the valve core (7) as the center and is connected to the inlet of the direct current flow channel, the width of the strip-shaped long opening (10) is equal to the inner diameter of the direct current flow channel, when the triangular cross-section outlet (8) gradually enters the direct current flow channel, the cross-sectional area of ​​the triangular cross-section outlet (8) in contact with the direct current flow channel gradually increases, and during the process of the valve core (7) rotating and opening in the valve body, the cross-sectional shape of the strip-shaped long opening (10) in contact with the direct current flow channel does not change with the rotation of the valve core (7); it is characterized in that The test method includes the following steps: Step 1: Install the flow control device inside the aircraft model, connect the inlet of the flow control device to the air supply pipeline, and connect the outlet of the flow control device to the ejector; Step 2: Measure the momentum coefficient C of the valve of the flow control device at different openings μ , the formula is as follows: In the formula, q ∞ is the dynamic pressure of the aircraft model during flight; S is the reference area of ​​the model; the gas mass flow rate in the pipeline Ejector outlet velocity u j Step 3: Conduct a static force test on the model in a wind tunnel to measure the effect of the jet on the six-element aerodynamic coefficient under different momentum coefficients; Step 4: According to the influence of momentum coefficient on aerodynamic coefficient under different valve openings, establish the rudder trim distribution strategy of coupled jet control. The specific trim formula is as follows: Outer aileron allocation formula: Inner aileron allocation formula: d e,right =d e,left =d q (3) Rudder and jet combination, distribution formula: Where, k1 is the rudder surface allocation ratio; k2 is the rudder surface-jet conversion ratio; δ a,left , δ a,right They are the left and right outer ailerons respectively; δ e,left , δ e,right They are the left and right inner ailerons respectively; δ wt,left , δ wt,right They are the left and right wing rudders respectively; M sl,left 、M sl,right are the left and right jet momentum respectively; δ p , δ q , δ r They are the actual control amount of roll, the actual control amount of pitch, and the actual control amount of yaw; Step 5: Design the control law based on the control surface allocation strategy. The control law architecture adopts the structure of inner and outer loops. The outer loop control is position loop control, which gives pitch angular velocity, yaw angular velocity and roll angle targets for angle of attack, sideslip angle and roll angle targets. The inner loop control is speed loop control, which gives deflection angle instructions for inner and outer ailerons and rudder for angular velocity targets. Both the inner and outer loops adopt feedback control methods. The outer loop control first obtains the angle of attack velocity target, sideslip angle velocity target and roll angle velocity target according to the angle of attack, sideslip angle and roll angle errors. The formula is as follows: Among them, α, α c are the angle of attack and the angle of attack target, respectively; β, β c are the sideslip angle and the sideslip angle target respectively; φ, φ c are roll angle and roll angle target respectively; are the angle of attack velocity target, sideslip angular velocity target, and roll angular velocity target, respectively. α , K β , K φ are the angle of attack control gain, sideslip angle control gain, and roll angle control gain respectively; The sideslip angular velocity is decomposed into the roll axis and yaw axis through the angle of attack, and the formula for obtaining the body axis angular velocity target is as follows: Among them, p c ,q c 、r c They are the roll axis angular velocity target, pitch axis angular velocity target, and yaw axis angular velocity target of the inner loop, which are used for inner loop control; The inner loop control performs rudder control according to the angular velocity target. The formula for the rudder control increment is as follows: Among them, p, q, and r are the angular velocity of the roll axis, the angular velocity of the pitch axis, and the angular velocity of the yaw axis, respectively. p , Δδ q , Δδ r are the roll control increment, pitch control increment, and yaw control increment, respectively, K p , K q , K r They are roll control gain, pitch control gain, and yaw control gain, respectively. The formula for adding the control increment and the control trim amount to obtain the actual control amount is as follows: Among them, δ p,0 , δ q,0 , δ r,0 They are roll trim, pitch trim, and yaw trim respectively; Step 6: Based on the control surface allocation strategy of step 4 and the control law of step 5, conduct a model flight test to verify the flight control capability of the ejector.

2. The method for testing a jet flow control device for an aircraft model flight test according to claim 1, characterized in that: The rotation angle range of the valve core (7) in the valve body is 0°-90°.

3. The method for testing a jet flow control device for an aircraft model flight test according to claim 1, characterized in that: The triangular cross-section outlet is a horizontally placed isosceles triangle, and the vertex angle (9) thereof is 30°-60°.

4. The method for testing a jet flow control device for an aircraft model flight test according to claim 1, characterized in that: The driving motor (1) is a brushless motor, the stall torque of which is greater than the valve core starting torque, the rotation angular velocity is not less than 300 degrees / second, and the signal frequency is 50 Hz.

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