Aircraft self-adaptive thrust steering control method based on attitude measurement

Abstract

The invention belongs to the technical field of aircraft control, and provides an aircraft self-adaptive thrust steering control method based on attitude measurement, which comprises the following steps: S1, measuring a pitch angle rate, a pitch angle and a speed pitch inclination angle of an aircraft; s2, solving a pitch angle error signal and a pitch angle error integral signal; s3, solving an error correlation item adaptive signal; s4, solving an angular rate related item adaptive signal; and S5, realizing steering control according to an aircraft thrust steering pitching master control signal. According to the method, an equivalent control item model is constructed through aerodynamic coefficients, all total signals are combined to form a thrust steering total control signal, the expected pitch angle of the aircraft is tracked and regulated, and the flight control parameters of the speed, the pitch inclination angle and the pitch angle are compensated in real time, so that the flight control precision is improved. Control gain parameters are adjusted to enable the aircraft to adapt to environment and aerodynamic parameter changes, so that an aircraft control system can dynamically compensate interference factors in real time, and the control precision and stability of the aircraft are improved.

Description

Technical Field

[0001] The invention belongs to the technical field of aircraft attitude control, and in particular relates to an aircraft adaptive thrust steering control method based on attitude measurement. Background Art

[0002] Aircraft, especially special-shaped aircraft, have attracted widespread attention due to their unique aerodynamic shape. Their control is also more difficult than conventional aircraft. In particular, the control of elliptical and disc-shaped aircraft often adopts redundant control configuration. On the one hand, it reduces the control difficulty; on the other hand, the redundant configuration can keep the entire system stable when a single control is unstable.

[0003] The control concept of existing aircraft is based on the simultaneous consideration of two dimensions, from the stability of the rotational torque to the stability of the vertical force of the pitch channel, which is somewhat complex. In addition, traditional aircraft generally use a composite control system of thrust steering and mass distance dual control system, one control system is responsible for the stable balance of force, and the other control system is responsible for the balance of torque. However, considering special circumstances, such as the failure of the mass distance control system, the thrust steering control alone can also support the basic stable flight of the aircraft. Therefore, it is necessary to propose an aircraft adaptive thrust steering control method based on attitude measurement. Summary of the invention

[0004] In view of the problems existing in the prior art, the present invention discloses an aircraft adaptive thrust steering control method based on attitude measurement, which measures the speed pitch inclination, pitch angle and pitch angle rate, and adaptively compensates for the uncertainties related to the speed pitch inclination, pitch angle, pitch angle rate and pitch angle error hinge according to the pitch angle error and the pitch angle rate, then constructs an equivalent control item through the model aerodynamic coefficient, and combines the pitch angle rate and the pitch angle error, the pitch angle error integral and the adaptive signal to form a total thrust steering control quantity, so as to realize the aircraft tracking and regulation of the desired pitch angle, and improve the control accuracy and stability of the aircraft.

[0005] The present invention provides an aircraft adaptive thrust steering control method based on attitude measurement, which comprises the following steps:

[0006] S1. Measure the pitch rate ω of the aircraft z , Pitch angle and velocity pitch angle θ;

[0007] S2. Install a thrust steering device at the axial tail of the aircraft in the flight direction. The thrust pitch swing angle of the thrust steering device is ξ. Set the desired pitch angle of the aircraft to The pitch angle ω of the aircraft z By comparison, the pitch angle error e is obtained, and the pitch angle error integral s1 of the aircraft is obtained by integration, which is specifically:

[0008]

[0009] S3, solving the error-related adaptive signal;

[0010] S4. Solve the adaptive signal of angular rate related items:

[0011] S5, solving the total control signal of the thrust steering pitch of the aircraft to achieve steering control, the sub-steps of which include:

[0012] S51, according to the pitch angle θ and the pitch angle of the aircraft As well as the aerodynamic parameters of the aircraft, design the thrust steering equivalent control signal f of the aircraft e :

[0013]

[0014] In the formula, a 24 、a z are the aerodynamic parameters of the aircraft, both of which are constants;

[0015] The expression of the dynamic control model in the pitch channel control of the aircraft is:

[0016] Δu=K1Δα+K1Δq

[0017] In the formula, Δu is the control input signal of the elevator, Δα is the pitch angle deviation, Δq is the pitch angle velocity deviation, and K1 and K1 are control gains;

[0018] Receive the pitch angle error e and pitch angle rate ω z After that, the velocity pitch angle θ and pitch angle Flight control parameters are compensated, and the aircraft is adapted to the environment and aerodynamic parameter changes by adjusting the control gain parameters, so that the aircraft control system can dynamically compensate for interference factors in real time;

[0019] S52, superimpose the pitch angle error e, the pitch angle error integral s1, and the pitch angle rate ω z , error-related adaptive signal f c The adaptive signal f related to the angular rate d , forming the total thrust steering pitch control signal of the aircraft, and transmitting it to the thrust steering device to control the thrust pitch swing angle of the engine to be equal to the total thrust steering pitch control signal of the aircraft, specifically:

[0020] ξ a =-l1ω z -l2e-l3s1+f c +f d -f e;

[0021] In the formula, ξ a is the total thrust steering pitch control signal; l1, l2, l3 are control parameters, all of which are constants;

[0022] The equivalent control term is constructed by using the model aerodynamic coefficients, and the total thrust steering control signal is composed of the pitch angle rate and pitch angle error, the pitch angle error integral, and the adaptive signal to achieve the desired pitch angle control of the aircraft.

[0023] Preferably, the step S3 includes the following sub-steps:

[0024] S31. According to the pitch angle error e and the velocity pitch angle θ, the adaptive growth rate c of the hinge factor of the error and the velocity pitch angle is calculated. 1d , and integrate to get the error and velocity pitch angle hinge adaptive signal c1;

[0025]

[0026] c1(n+1)=c1(n)+c 1d T

[0027] Where, k1 and k2 are constant parameters used to adjust the error and velocity pitch angle hinge adaptive signal growth rate, ε1 is a constant parameter used to soften the angle signal, and T is a constant integral parameter;

[0028] S32, according to the pitch angle error e and the pitch angle Solution error and pitch angle hinge factor adaptive growth rate c 2d , and integrate to get the error and pitch angle hinge adaptive signal c2:

[0029]

[0030] c2(n+1)=c2(n)+c 2d T

[0031] Where, k3 and k4 are constant parameters, which are used to adjust the error and pitch angle hinge adaptive signal growth rate;

[0032] S33, according to the pitch angle error e and the pitch angle rate ω z , solve the error and pitch angle rate hinge factor adaptive growth rate c 3d , and integrate to get the error and pitch angular rate hinge adaptive signal c3:

[0033]

[0034] c3(n+1)=c3(n)+c 3d T

[0035] Where, k5 and k6 are constant parameters, which are used to adjust the error and pitch angle rate hinge adaptive signal growth rate, and ε2 is a constant parameter;

[0036] S34. According to the pitch angle error e, solve the error factor adaptive growth rate c 4d , and integrate to obtain the error factor adaptive signal c4;

[0037]

[0038] c4(n+1)=c4(n)+c 4d T

[0039] Wherein, k7 and k8 are constant parameters used to adjust the error factor adaptive signal growth rate;

[0040] S35, signal superposition and comprehensive summary, to obtain the error-related adaptive signal f c :

[0041]

[0042] Preferably, the step S4 includes the following sub-steps:

[0043] S41, according to the pitch angle rate ω z and velocity pitch angle θ, solve the angular rate and velocity pitch angle hinge factor adaptive growth rate c 5d , and integrate to get the angular rate and velocity pitch angle hinge adaptive signal c5:

[0044]

[0045] c5(n+1)=c5(n)+c 5d T

[0046] Where: d1 and d2 are constant parameters, which are used to adjust the growth rate of the angular rate and velocity pitch angle hinge adaptive signal;

[0047] S42, according to the pitch angle rate ω z With pitch angle Solve the adaptive growth rate c of the angular rate and pitch angle hinge factor 6d , and integrate to get the angular rate and pitch angle hinge adaptive signal c6:

[0048]

[0049] c6(n+1)=c6(n)+c 6d T;

[0050] Where d3 and d4 are constant parameters, which are used to adjust the growth rate of the angular rate and pitch angle hinge adaptive signal;

[0051] S43, according to the pitch angle rate ω z , solve the adaptive growth rate c of the angular velocity factor 7d , and integrate to get the angular rate factor adaptive signal c7:

[0052]

[0053] c7(n+1)=c7(n)+c 7d T

[0054] Wherein, d5 and d6 are constant parameters, which are used to adjust the growth rate of the pitch angle rate adaptive signal;

[0055] S44, according to the pitch angle rate ω z and pitch angle error c 1d , solve the angular rate and pitch angle error hinge adaptive growth rate c 8d , and integrate to get the angular rate and pitch angle error hinge adaptive signal c8:

[0056]

[0057] c8(n+1)=c8(n)+c 8d T

[0058] Wherein, d7 and d8 are constant parameters, which are used to adjust the growth rate of the angular rate and pitch angle error hinge adaptive signal;

[0059] S45, signal superposition and comprehensive summary, to obtain the angular rate related adaptive signal f d :

[0060]

[0061] Preferably, firstly, by measuring the velocity pitch angle θ, the pitch angle and the pitch rate ω z , and then according to the pitch angle error e and the pitch angle rate ω z For velocity pitch angle θ, pitch angle θ, pitch angle rate ω z The pitch angle error e and the uncertainty related to the hinge are adaptively compensated to obtain the adaptive signals related to the angle and angular rate, respectively. Then, the equivalent control term is constructed through the model aerodynamic coefficient, and then the pitch angle rate ω is used to adjust the pitch angle. z The dynamic characteristics of the dynamics are designed by inversion and combined with the pitch angle rate ω z And the pitch angle error e, the pitch angle error integral s1, the adaptive signal f dIt constitutes the total thrust steering control amount, enabling the aircraft to track and control the desired pitch angle.

[0062] Preferably, by measuring the pitch rate ω of the aircraft z , Pitch angle The adaptive control of the aircraft thrust steering is realized by using the velocity pitch angle θ, specifically:

[0063] The error and velocity pitch angle hinge adaptive signal c1 is:

[0064] c1=∫(k1 e+k2 e)dt

[0065] The error and pitch angle hinge adaptive signal c2 are:

[0066] c2=∫(k3 e+k4θ)dt

[0067] The angular rate and velocity pitch angle hinge adaptive signal c5 is:

[0068] c5=∫(k5θ+k6θv)dt

[0069] Thrust steering pitch total control signal ξ a for:

[0070] ξ a =k7e+k8∫edt+k9e+adaptive signal

[0071] Wherein, k1, k2, k3, k4, k5 and k6 are adjustment parameters, and k7, k8 and k9 are control gain parameters.

[0072] Compared with the prior art, the present invention has the following advantages:

[0073] 1. The invention provides an aircraft adaptive thrust steering control method based on attitude measurement,

[0074] 1. According to the pitch angle error signal and the pitch angle rate signal, adaptive compensation is performed on various hinge-related interferences and uncertainties such as velocity pitch inclination, pitch angle, pitch angle rate and pitch angle error, and adaptive signals related to angle and angular rate are obtained respectively, so that the entire control system can adapt to changes in the environment and aerodynamic parameters.

[0075] 2. The aircraft adaptive thrust steering control method based on attitude measurement of the present invention constructs equivalent control items through model aerodynamic coefficients, reasonably and effectively utilizes the known structural information of the model, and the unknown part and the interference part can be adaptively compensated.

[0076] 3. The aircraft adaptive thrust steering control method based on attitude measurement of the present invention adopts the method of matching error integration, proportion and pitch angle rate, and does not need the assistance of mass distance control. When the mass distance control system fails, it can still independently ensure the stable flight of the aircraft.

[0077] 4. The aircraft adaptive thrust steering control method based on attitude measurement of the present invention does not need to consider the stability of the force in the vertical direction of pitch and the dynamic compensation and stability of the speed pitch angle, thereby making the entire system analysis and control logic and process simple and effective. BRIEF DESCRIPTION OF THE DRAWINGS

[0078] Figure 1 It is a flow chart of an aircraft adaptive thrust steering control method based on attitude measurement of the present invention;

[0079] Figure 2 A schematic diagram of a thrust steering device of the thrust steering control method of the present invention;

[0080] Figure 3 A graph showing a pitch rate signal of an aircraft according to an embodiment of a thrust steering control method of the present invention;

[0081] Figure 4 A graph showing the pitch angle of an aircraft according to an embodiment of the thrust steering control method of the present invention;

[0082] Figure 5 It is a graph of the aircraft speed pitch angle inclination according to an embodiment of the thrust steering control method of the present invention;

[0083] Figure 6 A graph showing an aircraft pitch angle error signal according to an embodiment of the thrust steering control method of the present invention;

[0084] Figure 7 A graph showing an integral signal of an aircraft pitch angle error according to an embodiment of a thrust steering control method of the present invention;

[0085] Figure 8 This is a graph of the total control signal of the aircraft thrust steering according to the thrust steering control method embodiment of the present invention. DETAILED DESCRIPTION

[0086] In order to fully describe the technical content, structural features, objectives and effects of the present invention, the following will be described in detail with reference to the accompanying drawings.

[0087] The invention provides an aircraft adaptive thrust steering control method based on attitude measurement, such as Figure 1 As shown, it includes the following steps:

[0088] S1. Measure the pitch rate ω of the aircraft z , Pitch angle The velocity pitch angle θ, its variation curve is as follows Figure 3 to Figure 5 shown.

[0089] like Figure 4 As shown, the aircraft pitch angle It rises to the expected value of 10° in about 1 second.

[0090] like Figure 5 As shown, the pitch angle error e converges to 0 in about 1s;

[0091] S2. Install a thrust steering device at the axial tail of the aircraft in the flight direction, such as Figure 2 As shown, this embodiment takes a circular aircraft as an example. The method provided by the present invention does not require the shape of the aircraft to be limited to a circle, a disc or an ellipse. The thrust pitch angle is ξ. According to the flight mission, the desired pitch angle of the aircraft is set to In this embodiment, the desired pitch angle signal of the aircraft is 10°, which is consistent with the pitch angle ω of the aircraft. z By comparison, we get the pitch angle error e. The pitch angle error e curve is as follows: Figure 6 As shown, the pitch angle error integral s1 is obtained by integration, which is specifically:

[0092]

[0093] S3, solving the error-related adaptive signal, which includes the following sub-steps:

[0094] S31. According to the pitch angle error e and the velocity pitch angle θ, the adaptive growth rate c of the hinge factor of the error and the velocity pitch angle is calculated. 1d , and integrate to get the error and velocity pitch angle hinge adaptive signal c1;

[0095]

[0096] c1(n+1)=c1(n)+c 1d T

[0097] Wherein, k1 and k2 are constant parameters used to adjust the error and the growth rate of the velocity pitch angle hinge adaptive signal. In this embodiment, k1=0.1, k2=0.15, ε1 is a constant parameter used to soften the angle signal, and T is a constant integral parameter;

[0098] S32, according to the pitch angle error e and the pitch angle Solution error and pitch angle hinge factor adaptive growth rate c 2d , and integrate to get the error and pitch angle hinge adaptive signal c2:

[0099]

[0100] c2(n+1)=c2(n)+c 2d T

[0101] Wherein, k3 and k4 are constant parameters used to adjust the error and pitch angle hinge adaptive signal growth rate. In this embodiment, k3=0.18, k4=0.25;

[0102] S33, according to the pitch angle error e and the pitch angle rate ω z , solve the error and pitch angle rate hinge factor adaptive growth rate c 3d , and integrate to get the error and pitch angular rate hinge adaptive signal c3:

[0103]

[0104] c3(n+1)=c3(n)+c 3d T

[0105] Wherein, k5 and k6 are constant parameters, which are used to adjust the error and the pitch angle rate hinge adaptive signal growth rate. In this embodiment, k5=0.35, k6=0.05, and ε2 is a constant parameter;

[0106] S34. According to the pitch angle error e, solve the error factor adaptive growth rate c 4d , and integrate to obtain the error factor adaptive signal c4;

[0107]

[0108] c4(n+1)=c4(n)+c 4d T

[0109] Wherein, k7 and k8 are constant parameters used to adjust the error factor adaptive signal growth rate, and in this embodiment, k7=0.03 and k8=0.08.

[0110] S35, signal superposition and comprehensive summary, to obtain the error-related adaptive signal f c :

[0111]

[0112] S4, solving the angular rate related adaptive signal, which includes the following sub-steps:

[0113] S41, according to the pitch angle rate ω z and velocity pitch angle θ, solve the angular rate and velocity pitch angle hinge factor adaptive growth rate c 5d , and integrate to get the angular rate and velocity pitch angle hinge adaptive signal c5:

[0114]

[0115] c5(n+1)=c5(n)+c 5d T

[0116] Wherein: d1 and d2 are constant parameters used to adjust the growth rate of the angular rate and the velocity pitch angle hinge adaptive signal. In this embodiment, d1=0.25, d2=0.13.

[0117] S42, according to the pitch angle rate ω z With pitch angle Solve the adaptive growth rate c of the angular rate and pitch angle hinge factor 6d , and integrate to get the angular rate and pitch angle hinge adaptive signal c6:

[0118]

[0119] c6(n+1)=c6(n)+c 6d T;

[0120] Wherein, d3 and d4 are constant parameters used to adjust the growth rate of the angular velocity and the pitch angle hinge adaptive signal. In this embodiment, d3=0.6 and d4=0.04 are selected.

[0121] S43, according to the pitch angle rate ω z , solve the adaptive growth rate c of the angular velocity factor 7d , and integrate to get the angular rate factor adaptive signal c7:

[0122]

[0123] c7(n+1)=c7(n)+c 7d T

[0124] Wherein, d5 and d6 are constant parameters used to adjust the growth rate of the pitch angle rate adaptive signal. In this embodiment, d5=0.33, d6=0.02.

[0125] S44, according to the pitch angle rate ω z and pitch angle error c 1d , solve the angular rate and pitch angle error hinge adaptive growth rate c 8d , and integrate to get the angular rate and pitch angle error hinge adaptive signal c8:

[0126]

[0127] c8(n+1)=c8(n)+c 8d T

[0128] Wherein, d7 and d8 are constant parameters used to adjust the growth rate of the angular velocity and the pitch angle error hinge adaptive signal. In this embodiment, d7=0.05, d8=0.1.

[0129] S45, signal superposition and comprehensive summary, to obtain the angular rate related adaptive signal f d :

[0130]

[0131] S5, solving the total control signal of the thrust steering pitch of the aircraft to achieve steering control, the sub-steps of which include:

[0132] S51, according to the pitch angle θ and the pitch angle of the aircraft As well as the aerodynamic parameters of the aircraft, design the thrust steering equivalent control signal f of the aircraft e :

[0133]

[0134] In the formula, a 24 、a z are the aerodynamic parameters of the aircraft, which are all constants. In this embodiment, a is obtained based on the wind tunnel test results of its external dimensions. 24 =829,a z =-12.89.

[0135] The expression of the dynamic control model in the pitch channel control of the aircraft is:

[0136] Δu=K1Δα+K1Δq

[0137] Wherein, Δu is the control input signal of the elevator, Δα is the pitch angle deviation, Δq is the pitch angle velocity deviation, and K1 and K2 are control gains.

[0138] In step S5, the pitch angle error e and the pitch angle rate ω are received. z After that, the velocity pitch angle θ and pitch angle The flight control parameters are compensated, and the gain parameters are adjusted to make the aircraft adapt to the environment and aerodynamic parameter changes, so that the aircraft control system can dynamically compensate for interference factors in real time.

[0139] S52, superimpose the pitch angle error e, the pitch angle error integral s1, and the pitch angle rate ω z , error-related adaptive signal f c The adaptive signal f related to the angular rate d , forming the total thrust steering pitch control signal of the aircraft, and transmitting it to the thrust steering device to control the thrust pitch swing angle of the engine to be equal to the total thrust steering pitch control signal of the aircraft, specifically:

[0140] ξ a =-l1ω z -l2e-l3s1+f c +f d -f e ;

[0141] In the formula, ξ a is the total thrust steering pitch control signal; l1, l2, l3 are control parameters, all of which are constants. In this embodiment, l1=0.5, l2=1.5, l3=0.1.

[0142] The thrust-steering pitch total control signal is obtained to enable the aircraft to track the desired pitch angle and complete the pitch channel control task.

[0143] By measuring the pitch rate ω of the aircraft z , Pitch angle The adaptive control of the aircraft thrust steering is realized by using the velocity pitch angle θ, specifically:

[0144] The error and velocity pitch angle hinge adaptive signal c1 is:

[0145] c1=∫(k1 e+k2 e)dt

[0146] The error and pitch angle hinge adaptive signal c2 are:

[0147] c2=∫(k3e+k4θ)dt

[0148] The angular rate and velocity pitch angle hinge adaptive signal c5 is:

[0149] c5=∫(k5θ+k6θv)dt

[0150] Thrust steering pitch total control signal ξ a for:

[0151] ξ a =k7e+k8∫edt+k9e+adaptive signal

[0152] Wherein, k1, k2, k3, k4, k5 and k6 are adjustment parameters, and k7, k8 and k9 are control gain parameters.

[0153] First, by measuring the velocity pitch angle θ, pitch angle and the pitch rate ω z , and then according to the pitch angle error e and the pitch angle rate ω z Pitch angle θ for speed, pitch angle Pitch rate ω zThe pitch angle error e and the uncertainty related to the hinge are adaptively compensated to obtain the adaptive signals related to the angle and angular rate, respectively. Then, the equivalent control term is constructed through the model aerodynamic coefficient, and then the pitch angle rate ω is used to adjust the pitch angle. z The dynamic characteristics of the dynamics are designed by inversion and combined with the pitch angle rate ω z And the pitch angle error e, the pitch angle error integral s1, the adaptive signal f d It constitutes the total thrust steering control amount, enables the aircraft to track the desired pitch angle, and completes the pitch channel control task.

[0154] like Figure 8 As shown in the figure, the total thrust steering pitch control signal is smooth and stable at 0, and the maximum value of the whole process does not exceed 4, and there is no obvious flutter and jitter. The control signal can meet the engineering requirements, and the response of the whole system has very good smoothness and rapidity.

[0155] The invention discloses an aircraft adaptive thrust steering control method based on attitude measurement, which can adaptively compensate for speed pitch inclination, pitch angle, angular rate and pitch angle error according to a pitch angle error signal and a pitch angle rate signal, construct an equivalent control item through aerodynamic coefficients of a model, and form a thrust steering total control signal by combining the pitch angle rate and the pitch angle error, the pitch angle error integral and the adaptive signal, so as to realize the regulation of the aircraft to a desired pitch angle.

[0156] Moreover, the present invention optimizes adaptive control by adjusting flight control parameters according to the specific characteristics of the aircraft and flight mission requirements to adapt to different flight requirements at each stage, and the control signal can accumulate error information, so as to gradually adjust the control strategy during long-term flight to adapt to the changing trend of the aircraft's attitude. When facing external interference, the integral term can continue to act to restore the aircraft to the desired flight state.

[0157] The embodiments described above are only descriptions of the preferred implementation modes of the present invention, and are not intended to limit the scope of the present invention. Without departing from the design spirit of the present invention, various modifications and improvements made to the technical solutions of the present invention by ordinary technicians in this field should all fall within the protection scope determined by the claims of the present invention.

Claims

1. An aircraft adaptive thrust steering control method based on attitude measurement, characterized in that: It includes the following steps: S1. Measure the pitch rate ω of the aircraft z , Pitch angle and velocity pitch angle θ; S2. Install a thrust steering device at the axial tail of the aircraft in the flight direction. The thrust pitch swing angle of the thrust steering device is ξ. Set the desired pitch angle of the aircraft to The pitch angle ω of the aircraft z By comparison, the pitch angle error e is obtained, and the pitch angle error integral s1 of the aircraft is obtained by integration, which is specifically: S3, solving the error-related adaptive signal; S4. Solve the adaptive signal of angular rate related items: S5, solving the total control signal of the thrust steering pitch of the aircraft to achieve steering control, the sub-steps of which include: S51, according to the pitch angle θ and the pitch angle of the aircraft As well as the aerodynamic parameters of the aircraft, design the thrust steering equivalent control signal f of the aircraft e : In the formula, a 24 、a z are the aerodynamic parameters of the aircraft, both of which are constants; The expression of the dynamic control model in the pitch channel control of the aircraft is: Δu=K1Δα+K1Δq In the formula, Δu is the control input signal of the elevator, Δα is the pitch angle deviation, Δq is the pitch angle velocity deviation, and K1 and K1 are control gains; Receive the pitch angle error e and pitch angle rate ω z After that, the velocity pitch angle θ and pitch angle Flight control parameters are compensated, and the aircraft is adapted to the environment and aerodynamic parameter changes by adjusting the control gain parameters, so that the aircraft control system can dynamically compensate for interference factors in real time; S52, superimpose the pitch angle error e, the pitch angle error integral s1, and the pitch angle rate ω z , error-related adaptive signal f c The adaptive signal f related to the angular rate d , forming the total thrust steering pitch control signal of the aircraft, and transmitting it to the thrust steering device to control the thrust pitch swing angle of the engine to be equal to the total thrust steering pitch control signal ξ of the aircraft a , specifically: x a =-l1ω z -l2e-l3s1+f c +f d -f e ; In the formula, l1, l2, l3 are control parameters, all of which are constants; The equivalent control term is constructed by using the model aerodynamic coefficients, and the total thrust steering control signal is composed of the pitch angle rate and pitch angle error, the pitch angle error integral, and the adaptive signal to achieve the desired pitch angle control of the aircraft.

2. The method for adaptive thrust steering control of an aircraft based on attitude measurement according to claim 1, characterized in that: The step S3 comprises the following sub-steps: S31. According to the pitch angle error e and the velocity pitch angle θ, solve the pitch angle error and the velocity pitch angle hinge factor adaptive growth rate c 1d , and integrate to get the error and velocity pitch angle hinge adaptive signal c1; c1(n+1)=c1(n)+c 1d T Where, k1 and k2 are constant parameters used to adjust the error and velocity pitch angle hinge adaptive signal growth rate, ε1 is a constant parameter used to soften the angle signal, and T is a constant integral parameter; S32, according to the pitch angle error e and the pitch angle Solution error and pitch angle hinge factor adaptive growth rate c 2d , and integrate to get the error and pitch angle hinge adaptive signal c2: c2(n+1)=c2(n)+c 2d T Where, k3 and k4 are constant parameters, which are used to adjust the error and pitch angle hinge adaptive signal growth rate; S33, according to the pitch angle error e and the pitch angle rate ω z , solve the error and pitch angle rate hinge factor adaptive growth rate c 3d , and integrate to get the error and pitch angular rate hinge adaptive signal c3: c3(n+1)=c3(n)+c 3d T Where, k5 and k6 are constant parameters, which are used to adjust the error and pitch angle rate hinge adaptive signal growth rate, and ε2 is a constant parameter; S34. According to the pitch angle error e, solve the error factor adaptive growth rate c 4d , and integrate to obtain the error factor adaptive signal c4; c4(n+1)=c4(n)+c 4d T Wherein, k7 and k8 are constant parameters used to adjust the error factor adaptive signal growth rate; S35, signal superposition and comprehensive summary, to obtain the error-related adaptive signal f c :

3. The method for adaptive thrust steering control of an aircraft based on attitude measurement according to claim 1, characterized in that: The step S4 comprises the following sub-steps: S41, according to the pitch angle rate ω z and velocity pitch angle θ, solve the angular rate and velocity pitch angle hinge factor adaptive growth rate c 5d , and integrate to get the angular rate and velocity pitch angle hinge adaptive signal c5: c5(n+1)=c5(n)+c 5d T Where: d1 and d2 are constant parameters, which are used to adjust the growth rate of the angular rate and velocity pitch angle hinge adaptive signal; S42, according to the pitch angle rate ω z With pitch angle Solve the adaptive growth rate c of the angular rate and pitch angle hinge factor 6d , and integrate to get the angular rate and pitch angle hinge adaptive signal c6: c6(n+1)=c6(n)+c 6d T; Where d3 and d4 are constant parameters, which are used to adjust the growth rate of the angular rate and pitch angle hinge adaptive signal; S43, according to the pitch angle rate ω z , solve the adaptive growth rate c of the angular velocity factor 7d , and integrate to get the angular rate factor adaptive signal c7: c7(n+1)=c7(n)+c 7d T Wherein, d5 and d6 are constant parameters, which are used to adjust the growth rate of the pitch angle rate adaptive signal; S44, according to the pitch angle rate ω z and pitch angle error c 1d , solve the angular rate and pitch angle error hinge adaptive growth rate c 8d , and integrate to get the angular rate and pitch angle error hinge adaptive signal c8: c8(n+1)=c8(n)+c 8d T Wherein, d7 and d8 are constant parameters, which are used to adjust the growth rate of the angular rate and pitch angle error hinge adaptive signal; S45, signal superposition and comprehensive summary, to obtain the angular rate related adaptive signal f d :

4. The method for adaptive thrust steering control of an aircraft based on attitude measurement according to claim 1, characterized in that: By measuring the pitch rate ω of the aircraft z , Pitch angle The adaptive control of the aircraft thrust steering is realized by using the velocity pitch angle θ, specifically: The error and velocity pitch angle hinge adaptive signal c1 is: c1=∫(k1e+k2e)dt The error and pitch angle hinge adaptive signal c2 are: c2=∫(k3e+k4θ)dt The angular rate and velocity pitch angle hinge adaptive signal c5 is: c5=∫(k5θ+k6θv)dt Thrust steering pitch total control signal ξ a for: ξ a =k7e+k8∫edt+k9e+adaptive signal Wherein, k1, k2, k3, k4, k5 and k6 are adjustment parameters, and k7, k8 and k9 are control gain parameters.

5. The method for adaptive thrust steering control of an aircraft based on attitude measurement according to claim 4, characterized in that: By measuring the velocity pitch angle θ, pitch angle and the pitch rate ω z , and then according to the pitch angle error e and the pitch angle rate ω z Pitch angle θ for speed, pitch angle Pitch rate ω z The pitch angle error e and the uncertainty related to the hinge are adaptively compensated to obtain the adaptive signals related to the angle and angular rate, respectively. Then, the equivalent control term is constructed through the model aerodynamic coefficient, and then the pitch angle rate ω is used to adjust the pitch angle. z The dynamic characteristics of the dynamics are designed by inversion and combined with the pitch angle rate ω z And the pitch angle error e, the pitch angle error integral s1, the adaptive signal f d It constitutes the total thrust steering control amount, enabling the aircraft to track and control the desired pitch angle.

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