Stability augmentation flight control method based on distributed pressure measurement and distributed propulsion

Abstract

The invention belongs to the technical field of flight control, and discloses a stability augmentation flight control method based on distributed pressure measurement and distributed propulsion. The stability augmentation flight control method comprises the steps that the distributed propellers and the distributed pressure sensors are installed; pneumatic data and pressure distribution under the differential power are obtained; measuring wing surface pressure distribution; establishing a distributed propeller differential control law; differential control of the distributed propellers is carried out; and the course stability and the attitude stability are improved. According to the stability augmentation flight control method, propellers and pressure sensors are distributed on the upper wing surfaces of wings on the two sides of an aircraft, pressure distribution on the two sides of the wings is measured, the distributed propellers are controlled in a differential mode, flow fields on the two sides of the wings are adjusted, lift augmentation and drag reduction are conducted, and therefore power differential wind-resistant flight control driven by flow field perception is achieved. The course stability and attitude stability of the aircraft are improved, and the flight safety of the aircraft in complex wind field environments such as strong wind power and rapid wind direction change is improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of flight control, and particularly relates to a stability augmentation flight control method based on distributed pressure measurement and distributed propulsion. Background Art

[0002] Distributed power aircraft is a new layout and new concept aircraft, and its propulsion system consists of multiple electric propellers or small ducted fans distributed on the fuselage or wings. The distributed power method can reduce the requirements and design difficulties of a single large engine, reduce aircraft noise, improve stability and wind disturbance resistance, improve the safety of the propulsion system, and shorten the takeoff and landing distance, and has wide applications in both civilian and military fields. However, the new layout and new propulsion method also face many problems. For example, during the power differential stability augmentation and wind resistance flight process, how to determine the thrust that each ducted fan needs to provide, and what is the basis for stability augmentation control. If this problem is not solved, it is difficult to realize the power differential stability augmentation and wind resistance flight technology.

[0003] The existing stability augmentation flight control methods mainly achieve it by aircraft variant or changing control parameters. The Chinese patent literature database discloses an invention named a control method for a variable stability stealth aircraft based on an embedded vertical tail (CN115857523A). This control method improves the yaw stability and maneuverability of a tailless aircraft during large maneuvers by designing the control law of the embedded control surface with respect to the aircraft bank angle. Although this control method can enhance the yaw stability of the aircraft, however, due to the need for variant, the stability and failure rate during the variant transition process cannot be ignored.

[0004] The Chinese patent literature database discloses an invention named a yaw moment suppression method for a distributed electric propulsion aircraft (CN117416221A). This yaw moment suppression method obtains the yaw moment generated by internal faults of the distributed electric propulsion system according to the motor state parameters and aircraft characteristic parameters, and then suppresses the yaw moment caused by faults by controlling the speed of each motor. This yaw moment suppression method monitors the motor state, converts the corresponding faults into yaw moments, and enhances the yaw stability after faults. However, this yaw moment suppression method cannot obtain the influence of environmental factors such as wind disturbance, and cannot well offset the yaw moment when the wind is strong.

[0005] Currently, there is an urgent need to develop a stability augmentation flight control method based on distributed pressure measurement and distributed propulsion. Summary of the Invention

[0006] The technical problem to be solved by the present invention is to provide a stability augmentation flight control method based on distributed pressure measurement and distributed propulsion to overcome the defects of the prior art.

[0007] During actual flight, due to the influence of environmental wind disturbances, the true airspeed is not equal to the ground speed, and the oncoming flow direction is not directly facing the nose of the aircraft. To counteract the influence of environmental wind disturbances on course stability, the enhanced stability flight control method based on distributed pressure measurement and distributed propulsion of the present invention measures the pressure distribution on the wing surface through distributed pressure sensors, and differentially controls the distributed thrusters according to the pressure distribution, so as to reduce the flow field difference between the two wings and improve the course stability and attitude stability of the aircraft.

[0008] The enhanced stability flight control method based on distributed pressure measurement and distributed propulsion of the present invention includes the following steps:

[0009] S10. Install distributed thrusters and distributed pressure sensors;

[0010] Install distributed thrusters and distributed pressure sensors on the upper wing surfaces of the two wings of the aircraft; the thrusters are installed at the trailing edges of the upper wing surfaces and are symmetrically distributed left and right; the pressure sensors are installed on the surfaces of the upper wing surfaces and are symmetrically distributed left and right;

[0011] Among them, Y ij represents a group of pressure sensors arranged at the i-th row and j-th column on both sides of the wing, Y ij左 represents the pressure sensor arranged at the i-th row and j-th column on the left wing, Y ij右 represents the pressure sensor arranged at the i-th row and j-th column on the right wing, Y represents the asymmetric resultant force on both sides of the wing; L ij represents the distance between the pressure sensor arranged at the i-th row and j-th column on each wing and the longitudinal axis of the fuselage, M ij represents the asymmetric moment generated by a group of pressure sensors arranged at the i-th row and j-th column on both sides of the wing, M represents the asymmetric resultant moment on both sides of the wing; m represents the number of rows of pressure sensors on each side, and n represents the number of columns of pressure sensors on each side;

[0012] The calculation formulas are as follows:

[0013]

[0014]

[0015] S20. Obtain aerodynamic data and pressure distribution under differential power;

[0016] According to the differential power of the thrusters on the two wings of the aircraft set in advance, conduct a wind tunnel test or numerical simulation to obtain the aerodynamic data and pressure distribution under differential power, and obtain the roll moment curve and yaw moment curve under the differential control of the thrusters;

[0017] S30. Measure the pressure distribution on the wing surface;

[0018] Conduct a flight test, and calculate the asymmetric forces and asymmetric moments on both sides of the aircraft wing caused by wind disturbances based on the pressure distribution measured by the distributed pressure sensors;

[0019] S40. Establish a differential control law for the distributed thrusters;

[0020] Divide the asymmetric forces and asymmetric moments on both sides of the wing into four modes, namely: Y k ∈{Y 1 ,Y 2 ,Y 3 ,Y 4} and M h ∈{M 1 ,M 2 ,M 3 ,M 4}, where k is the mode of the asymmetric force on both sides of the wing, and the value of k is 1, 2, 3, or 4; h is the mode of the asymmetric moment on both sides of the wing, and the value of h is 1, 2, 3, or 4;

[0021] Establish a differential control law D for the distributed thrusters:

[0022]

[0023] Among them, no differential means that the thrusters on both sides of the wing are fully opened at the maximum power; 20% differential means that the thruster on the left side of the wing is at 100% power and the thruster on the right side is at 80% power; 50% differential means that the thruster on the left side of the wing is at 100% power and the thruster on the right side is at 50% power; 100% differential means that the thruster on the left side of the wing is at 100% power and the thruster on the right side is idling;

[0024] S50. Conduct differential control of the distributed thrusters;

[0025] Judge which mode in the differential control law D of the distributed thrusters in S40 the asymmetric force Y k and the asymmetric moment M h on both sides of the wing are in during the flight test;

[0026] S60. Improve the course stability and attitude stability;

[0027] Conduct differential control of the distributed thrusters according to Equation (3) in S40 to improve the course stability and attitude stability.

[0028] Furthermore, the thruster is a ducted fan.

[0029] The flight control method for stability augmentation based on distributed pressure measurement and distributed propulsion of the present invention distributes thrusters and pressure sensors on the upper wing surfaces of both wings of the aircraft, measures the pressure distribution on both sides of the wings, and controls the distributed thrusters differentially to narrow the flow fields on both sides of the wings, increasing lift and reducing drag, thereby realizing the dynamic differential anti-wind flight control driven by flow field perception, improving the heading stability and attitude stability of the aircraft, and enhancing the flight safety of the aircraft in complex wind field environments such as strong wind and rapid wind direction change. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] Figure 1 It is a schematic structural diagram of a distributed power aircraft;

[0031] Figure 2 It is a curve of the influence of dynamic differential of a distributed power aircraft on yaw moment;

[0032] Figure 3 It is a curve of the influence of dynamic differential of a distributed power aircraft on roll moment;

[0033] Figure 4 It is a flowchart of the flight control method for stability augmentation based on distributed pressure measurement and distributed propulsion of the present invention.

[0034] In the figure, 1. Thruster; 2. Pressure sensor. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0035] The present invention will be described in detail below with reference to the drawings and embodiments.

[0036] Embodiment: The distributed power aircraft of this embodiment is shown in Figure 1 , Figure 1 where the label 1 represents a thruster and the label 2 represents a pressure sensor.

[0037] As Figure 4 shown, the flight control method for stability augmentation based on distributed pressure measurement and distributed propulsion of this embodiment includes the following steps:

[0038] S10. Install distributed thrusters and distributed pressure sensors;

[0039] Install distributed thrusters and distributed pressure sensors on the upper wing surfaces of both wings of the aircraft; the thrusters are installed at the trailing edges of the upper wing surfaces and are symmetrically distributed left and right; the pressure sensors are installed on the surfaces of the upper wing surfaces and are symmetrically distributed left and right;

[0040] Among them, Y ij represents a group of pressure sensors arranged in the i-th row and j-th column on both sides of the wing, and Y ij左 represents the pressure sensor arranged in the i-th row and j-th column on the left wing, and Y ij右denotes the pressure sensor arranged at the \(i\)-th row and \(j\)-th column of the right wing, \(Y\) denotes the asymmetric resultant force on both sides of the wing; \(L\) ij denotes the distance between the pressure sensor arranged at the \(i\)-th row and \(j\)-th column of each wing and the longitudinal axis of the fuselage, \(M\) ij denotes the asymmetric moment generated by a group of pressure sensors arranged at the \(i\)-th row and \(j\)-th column on both sides of the wing, \(M\) denotes the asymmetric resultant moment of both sides of the wing; \(m\) denotes the number of rows of pressure sensors on each side, \(n\) denotes the number of columns of pressure sensors on each side;

[0041] The calculation formula is as follows:

[0042]

[0043]

[0044] S20. Obtain the aerodynamic data and pressure distribution under differential power;

[0045] According to the differential power of the thrusters on both sides of the aircraft set in advance, conduct a wind tunnel test or numerical simulation to obtain the aerodynamic data and pressure distribution under differential power, and obtain the roll moment curve and yaw moment curve under differential control of the thrusters as shown in Figure 2 、 Figure 3 ;

[0046] S30. Measure the pressure distribution on the wing surface;

[0047] Conduct a flight test, and calculate the asymmetric force and asymmetric moment on both sides of the aircraft wing caused by wind disturbance through the pressure distribution measured by the distributed pressure sensors;

[0048] S40. Establish a distributed thruster differential control law;

[0049] Divide the asymmetric force and asymmetric moment on both sides of the wing into four modes, namely: \(Y\) k \(\in\{Y\) 1 ,Y 2 ,Y 3 ,Y 4 \}\) and \(M\) h \(\in\{M\) 1 ,M 2 ,M 3 ,M 4 \}, where \(k\) is the mode of the asymmetric force on both sides of the wing, and the value of \(k\) is 1, 2, 3, or 4; \(h\) is the mode of the asymmetric moment on both sides of the wing, and the value of \(h\) is 1, 2, 3, or 4;

[0050] Establish a distributed thruster differential control law \(D\):

[0051]

[0052] Among them, no differential means that the thrusters on both sides of the wing are fully open at the maximum power; 20% differential means that the thruster on the left side of the wing is at 100% power and the thruster on the right side is at 80% power; 50% differential means that the thruster on the left side of the wing is at 100% power and the thruster on the right side is at 50% power; 100% differential means that the thruster on the left side of the wing is at 100% power and the thruster on the right side is at idle speed.

[0053] S50. Perform differential control of the distributed thrusters;

[0054] Judge which mode in the distributed thruster differential control law D of S40 the asymmetric force Y k and the asymmetric moment M h on both sides of the wing are in during the flight test;

[0055] S60. Improve the course stability and attitude stability;

[0056] Perform differential control on the distributed thrusters according to Equation (3) of S40 to improve the course stability and attitude stability.

[0057] Furthermore, the thruster is a ducted fan.

[0058] Although the embodiments of the present invention have been disclosed as above, they are not limited to the applications listed in the specification and embodiments. For those familiar with the art, without departing from the principle of the present invention, all the features disclosed in the present invention, or all the steps in the disclosed methods or processes, except for the mutually exclusive features and / or steps, can be combined in any way. The present invention is not limited to the specific details and the illustrated examples here.

Claims

1. A stabilized flight control method based on distributed pressure measurement and distributed propulsion, characterized in that: The stabilized flight control method comprises the following steps: S10. Install distributed thrusters and distributed pressure sensors; Distributed propellers and distributed pressure sensors are installed on the upper wing surfaces of both wings of the aircraft; the propellers are installed on the trailing edge of the upper wing surface and are symmetrically distributed on the left and right; the pressure sensors are installed on the surface of the upper wing surface and are symmetrically distributed on the left and right; Among them, Y iiii represents a group of pressure sensors arranged in the i-th row and j-th column on both sides of the wing, Y iiii左 represents the pressure sensor arranged in the i-th row and j-th column of the left wing, Y iiii右 represents the pressure sensor arranged in the i-th row and j-th column of the right wing, Y represents the asymmetric force on both sides of the wing; L iiii M represents the distance between the pressure sensor arranged in the i-th row and j-th column of each wing and the longitudinal axis of the fuselage. iiii represents the asymmetric torque generated by a group of pressure sensors arranged in the i-th row and j-th column on both sides of the wing, M represents the asymmetric resultant torque on both sides of the wing; m represents the number of rows of pressure sensors on each side, and n represents the number of columns of pressure sensors on each side; The calculation formula is as follows: S20. Obtaining aerodynamic data and pressure distribution under differential power; According to the preset differential power of the wing thrusters on both sides of the aircraft, a wind tunnel test or numerical simulation is performed to obtain aerodynamic data and pressure distribution under differential power, and to obtain a rolling moment curve and a yaw moment curve under thruster differential control; S30. Measure the pressure distribution on the wing surface; Conduct flight tests to calculate the asymmetric forces and moments on both sides of the aircraft wings caused by wind disturbances using the pressure distribution measured by the distributed pressure sensor; S40. Establish distributed thruster differential control law; The asymmetric forces and moments on both sides of the wing are divided into four modes: kk ∈{Y1,Y2,Y3,Y4} and M h ∈{M1,M2,M3,M4}, where k is the mode of the asymmetric force on both sides of the wing, and the value of k is 1, 2, 3 or 4; h is the mode of the asymmetric moment on both sides of the wing, and the value of h is 1, 2, 3 or 4; Establish the distributed thruster differential control law D: Among them, no differential means that the propellers on both sides of the wing are fully opened at maximum power; 20% differential means that the propeller on the left side of the wing is 100% power and the propeller on the right side is 80% power; 50% differential means that the propeller on the left side of the wing is 100% power and the propeller on the right side is 50% power; 100% differential means that the propeller on the left side of the wing is 100% power and the propeller on the right side is idling; S50. Perform distributed thruster differential control; Determine the asymmetric force Y on both sides of the wing during the flight test kk and asymmetric moment M h Which mode in the distributed thruster differential control law D at S40? S60. Improved heading stability and attitude stability; According to equation (3) of S40, the distributed thrusters are differentially controlled to improve the heading stability and attitude stability.

2. The method for stabilization flight control based on distributed pressure measurement and distributed propulsion according to claim 1, characterized in that: The propeller is a ducted fan.

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