A tracking control method for cruise phase of hypersonic vehicle with plasma excitation device

Abstract

The invention relates to the field of aircraft guidance control, and discloses a cruise tracking control method for a hypersonic aircraft with a plasma excitation device. The method comprises the following steps: step S1, establishing a dynamic model of a longitudinal motion plane of the aircraft; step S2, control-oriented longitudinal motion plane model conversion; step S3, establishing an equivalent control surface model of the plasma excitation device; step S4, designing a nonlinear disturbance observer; step S5, designing a speed control law based on the disturbance observer; step S6, designing an altitude control law based on the disturbance observer; and step S7, validating the effectiveness of a control algorithm. The invention additionally considers a plasma excitation device as an actuator, effectively alleviating the problem of low control efficiency of the control surface of the hypersonic aircraft in large airspace flight, and combining the nonlinear inverse control theory and the disturbance observer method, realizes high-precision trajectory tracking control of the aircraft with the plasma excitation device in the cruise stage.

Description

Technical Field

[0001] The invention relates to the field of aircraft guidance control, in particular to a cruise tracking control method for a hypersonic aircraft with a plasma excitation device. Background Art

[0002] When a hypersonic vehicle (flying at a speed greater than 5 times the Mach number) maneuvers in near-space, the surface of the vehicle and the external atmospheric environment undergo complex aerodynamic coupling heating effects. The fluid molecules close to the surface of the vehicle undergo complex physical and chemical reactions such as ionization and dissociation under the action of high temperature, forming a non-uniform plasma layer wrapped around the surface of the vehicle, the so-called "plasma sheath". The plasma sheath significantly changes the physical and chemical properties of the flight environment, posing a huge challenge to flight safety. Due to the special electromagnetic medium characteristics of the plasma sheath, electromagnetic signals will be cut off, absorbed, reflected, refracted, etc., causing microwave communications to not work properly when the aircraft flies through the atmosphere at high speed, thereby affecting the navigation and communication of the aircraft. When flying in near-space, the reduction in the efficiency of the aircraft's control surfaces also poses a huge challenge to flight safety. During the flight of hypersonic vehicles, processes such as thermochemical non-equilibrium, energy excitation, radiation, ablation, and coupling between energy modes occur. In these processes, the entire plasma has multi-field coupling effects such as high temperature, hypersonic flow field, multi-component gas mixing field, thermochemical non-equilibrium field, different energy mode fields, diffusion, and heat conduction processes.

[0003] In order to improve the safety and reliability of hypersonic vehicles in near-space flight and reduce the impact of plasma sheath on the aircraft, as early as the 1990s, researchers from various countries conducted research on magnetohydrodynamic (MHD) control technology. The relevant technology has been applied in the fields of magnetohydrodynamic power generation, magnetohydrodynamic energy diversion, and hypersonic vehicle inlet flow control. A magnetic field generator is embedded inside the aircraft, and the mutual interference between the magnetic field and the external fluid is controlled to achieve flow control on the surface of the aircraft, thereby reducing the surface resistance of the aircraft and improving the maneuverability. The plasma excitation device is another device that can generate plasma and is arranged on the surface of the aircraft. By applying a certain amount of energy excitation, a large number of charged ions are released to the external flight environment. The charged ions and the external high-temperature gas work together to change the external flow field structure of the aircraft, optimize the aerodynamic lift and drag characteristics of the aircraft, thereby solving the problem of reduced efficiency of the control surface of the hypersonic aircraft and effectively improving the maneuverability and flight safety of the hypersonic aircraft. Although the performance of aircraft can be improved through certain magnetofluid control technologies, considering that the conventional actuators of aircraft are pneumatic control surfaces and aircraft engines, how to efficiently integrate plasma excitation devices with conventional actuators and then give full play to the performance advantages of each actuator is one of the key scientific problems that need to be solved at present. Based on this, the present invention proposes a cruise segment tracking control method for a hypersonic aircraft with a plasma excitation device to solve the above problem. Summary of the invention

[0004] The object of the present invention is to provide a cruise phase tracking control method for a hypersonic vehicle with a plasma excitation device to solve the problems raised in the above-mentioned background technology.

[0005] To achieve the above object, the present invention provides the following technical solution: a cruise tracking control method for a hypersonic vehicle with a plasma excitation device, comprising the following steps:

[0006] Step S1, establishing a longitudinal motion plane dynamics model of the aircraft;

[0007] Step S2, control-oriented longitudinal motion plane model conversion;

[0008] Step S3, establishing an equivalent control surface model of the plasma excitation device;

[0009] Step S4, design of nonlinear disturbance observer;

[0010] Step S5, designing a speed control law based on a disturbance observer;

[0011] Step S6, designing a height control law based on a disturbance observer;

[0012] Step S7, verifying the effectiveness of the control algorithm.

[0013] Preferably: Step S1 is based on the external force on the aircraft and the rotation relationship around the center of mass, and the longitudinal motion plane model of the hypersonic aircraft is as follows:

[0014]

[0015] where m, h, V, α, θ, q and J y They are respectively the mass of the aircraft, flight altitude, speed, angle of attack, trajectory inclination, pitch rate and moment of inertia around the Y axis of the missile body, T, L, D, M t They are engine thrust, aerodynamic lift, drag and pitch moment respectively. The specific expressions are:

[0016]

[0017] Among them C T ,C L ,C D ,C m are the engine thrust coefficient, lift coefficient, drag coefficient and pitch moment coefficient, and their values ​​are related to the engine valve opening δ T ′, flight attack angle α, pitch control surface deflection angle δ z 'related, They are dynamic pressure, engine thrust reference area, aerodynamic force reference area, distance from engine thrust to vehicle moment reference point and aerodynamic moment length.

[0018] Preferably, the hypersonic vehicle longitudinal motion plane model after the control-oriented longitudinal motion plane model conversion in step S2 is:

[0019]

[0020] Preferably, the lift coefficient increment of the plasma excitation device in step S3 has the following relationship with the power P and the number N of the plasma excitation device:

[0021]

[0022] Where ΔC L is the lift coefficient increment, in %, N is the number of excitation device groups, P is the excitation power, and under the condition of the presence of plasma excitation device, the thrust, lift and torque coefficients of the aircraft (4) are modified as follows:

[0023]

[0024] in is the distance between the plasma excitation device and the moment reference point of the aircraft, and the equivalent rudder deflection is defined as follows:

[0025]

[0026] In the presence of a plasma excitation device, the hypersonic vehicle longitudinal motion plane model (13) is modified to:

[0027]

[0028] The definitions of other variables in model (17) are the same as those in model (13).

[0029] Preferably: the nonlinear disturbance observer design in step S4 considers the following uncertain dynamic system:

[0030]

[0031] Where v(t)∈R is the system state variable, F(t), G(t) are continuous functions, d(t) is the system uncertainty disturbance, and u(t) is the system input variable. For the uncertain system (21), the designed nonlinear disturbance observer is as follows:

[0032]

[0033] in are the estimated values ​​of the system states v(t) and d(t), respectively. If R→+∞, then

[0034]

[0035] Preferably: the speed control law of step S5 is designed as follows:

[0036]

[0037] Among them, k v1 ,k v2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d v The estimated value of

[0038]

[0039] Among them, R V ,a v1 ,a v2 ∈R + are the parameters to be designed, is the estimated value of V.

[0040] Preferably, the altitude control law in step S6 is designed as the altitude subsystem designs two intermediate virtual control laws and the final actual control law to achieve altitude maintenance and tracking, wherein one virtual control law θ is designed as:

[0041]

[0042] Among them, k γ1 ,k γ2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d γ The estimated value of

[0043]

[0044] Among them, R γ ,a γ1 ,a γ2 ∈R + are the parameters to be designed, is the estimated value of γ, and another virtual control law Q is designed as,

[0045]

[0046] Among them, k θ1 ,k θ2 are the parameters to be designed, and the actual control law is designed as follows:

[0047]

[0048] Among them, k Q1 ,k Q2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d Q The estimated value of

[0049]

[0050] Among them, R Q ,a Q1 ,a Q2 ∈R + are the parameters to be designed, is the estimated value of Q.

[0051] Compared with the prior art, the present invention has the following beneficial effects:

[0052] 1) Under given plasma excitation power and number of groups, equivalent conversion from plasma excitation device to aerodynamic control surface can be achieved;

[0053] 2) Compared with the conventional aircraft dynamics model, the present invention additionally considers the plasma excitation device as an actuator, which effectively alleviates the problems of low control efficiency of the control surface of the hypersonic aircraft in large airspace flight;

[0054] 3) By combining nonlinear inverse control theory and disturbance observer method, high-precision trajectory tracking control of the aircraft with plasma excitation device in the cruise phase is achieved. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] Figure 1 is a flow chart of the present invention;

[0056] Figure 2 This is a flow field diagram of an aircraft with a plasma excitation device;

[0057] Figure 3 It is the graph of plasma excitation device power / number of groups and lift coefficient increment;

[0058] Figure 4 It is a simulation diagram of the speed tracking of a hypersonic aircraft;

[0059] Figure 5 It is a simulation diagram of the altitude tracking of a hypersonic aircraft;

[0060] Figure 6 It is a simulation diagram of the ballistic inclination tracking of a hypersonic aircraft;

[0061] Figure 7 It is a simulation diagram of the pitch angle tracking of the hypersonic aircraft attitude;

[0062] Figure 8 This is a simulation diagram of the pitch rate tracking of a hypersonic aircraft;

[0063] Fig. 9 It is a diagram of valve opening of a hypersonic vehicle engine;

[0064] Fig.10 This is a diagram of power variation of plasma excitation device of hypersonic vehicle;

[0065] Fig.11 It is a diagram of the change of the deflection angle of the control surface of a hypersonic aircraft;

[0066] Fig.12 It is the estimation diagram of NDO interference to the velocity channel;

[0067] Fig.13 This is the estimation diagram of NDO interference with the ballistic inclination channel. DETAILED DESCRIPTION

[0068] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0069] Example

[0070] See also Figure 1-Figure 3 , a tracking control method for a cruise phase of a hypersonic vehicle with a plasma excitation device shown in the figure comprises the following steps:

[0071] Step S1, establish the aircraft longitudinal motion plane dynamics model, first give the coordinate system definition used in the present invention, including

[0072] Geocentric inertial coordinate system (o I x I y I z I ): origin of coordinate system o I It is the center of the earth. I z I The axis is perpendicular to the Earth's equatorial plane and points to the North Pole. I x I Axis and o I y I The axis is in the equatorial plane, o I x I It is the intersection line of the equatorial plane and the meridian plane, o I y I The axis is determined by the right-hand rule.

[0073] Ground coordinate system (ox g y g z g ): Consider the earth as a plane, select a point on the earth's surface as the origin o, and z g Axis plumb downward; x g In the horizontal plane, its direction can be arbitrarily chosen and is usually related to the flight mission. g The axis is determined according to the right-hand rule.

[0074] Body coordinate system (ox b y b z b ): The origin of the coordinate system is the center of mass of the aircraft, ox b The axis coincides with the longitudinal axis of the aircraft body, and the direction pointing to the head is positive, oz b Located in the longitudinal symmetry plane of the aircraft body and ox b The axis is vertical, pointing upward is positive, and oy b It is determined by the right-hand rule.

[0075] Velocity coordinate system (ox a y a z a ): The coordinate origin o is the center of mass of the aircraft, ox a Axis along the direction of the aircraft velocity, oz a The axis is located in the main symmetry plane of the aircraft and is perpendicular to ox a The axis points upward, oy a Axis and ox a z a The faces are perpendicular and form a right-handed coordinate system.

[0076] The flight envelope of a hypersonic vehicle can range from 120km to 30km, and the flight envelope spans a large range. Therefore, when studying its motion law, the gravitational acceleration can no longer be simply regarded as a constant, but should be described as a function of altitude. If the earth is regarded as a homogeneous sphere, the gravitational acceleration satisfies the following relationship:

[0077]

[0078] Among them, g 0 =9.81m / s 2 is the gravitational acceleration at sea level, R e =6378km is the radius of the earth.

[0079] The present invention adopts the 1976 American Standard Atmospheric Model to describe atmospheric parameters. The main parameters include altitude, temperature, speed of sound, pressure and air density. It is considered that temperature T, speed of sound c, pressure P and air density ρ are all functions of altitude and have the following relationship:

[0080]

[0081] Among them, ρ 0 =1.2258kg / m 3 ,β 0 =1.3785×10 -4 .

[0082] By analyzing the external forces on the aircraft and the relationship between the rotation around the center of mass, the following longitudinal motion plane model of the hypersonic aircraft can be obtained:

[0083]

[0084] where m, h, V, α, θ, q and J y They are respectively the mass of the aircraft, flight altitude, speed, angle of attack, trajectory inclination, pitch rate and moment of inertia around the Y axis of the missile body, T, L, D, M t They are engine thrust, aerodynamic lift, drag and pitch moment respectively. The specific expressions are:

[0085]

[0086] Among them C T ,C L ,C D ,C m are the engine thrust coefficient, lift coefficient, drag coefficient and pitch moment coefficient, and their values ​​are related to the engine valve opening δ T ′, flight attack angle α, pitch control surface deflection angle δ z 'related, They are dynamic pressure, engine thrust reference area, aerodynamic force reference area, distance from engine thrust to vehicle moment reference point, and aerodynamic moment length;

[0087] Step S2, control-oriented longitudinal motion plane model conversion. The working environment of hypersonic aircraft is complex and there are large modeling and aerodynamic errors. Therefore, the actual aircraft dynamics model belongs to a strongly uncertain system. In order to complete the flight mission efficiently, it is necessary to design a robust controller that can resist strong uncertainty systems.

[0088] In order to facilitate the design of robust controller, the original nonlinear model needs to be transformed. For the speed subsystem, the control target is to design a robust backstepping controller to make the flight speed V stably track V under the condition of strong external uncertainties. ref , the speed subsystem is,

[0089]

[0090] Substituting aerodynamic force, torque, and engine thrust (4) into equation (5), we can obtain:

[0091]

[0092] in

[0093]

[0094] The control objective of the height subsystem is: Under the conditions of strong external disturbance and model uncertainty, the backstepping control strategy is adopted to make h stably track h by designing a series of virtual control laws and the final actual control law. ref , the height error is defined as

[0095]

[0096] The track angle command is designed as

[0097]

[0098] Among them, k,k I is the parameter to be designed. If γ approaches γ d , the dynamic response is

[0099]

[0100] Under zero initial conditions, Laplace transform is performed on the above equation, and its characteristic equation is:

[0101] s 2 +ks+k I =0(10)

[0102] The characteristic roots of the above formula are and are all negative real numbers, so the system represented by equation (10) is stable. exponential convergence, so we only need to ensure that γ tracks γ d The height can effectively track the reference input. The rest of the height subsystem is

[0103]

[0104] In order to facilitate the design of the backstepping control law, the rest of the height subsystem is rewritten as follows:

[0105]

[0106] in

[0107]

[0108]

[0109]

[0110] Through the above transformation, the control-oriented hypersonic vehicle longitudinal motion plane model is:

[0111]

[0112] Step S3, establishing an equivalent control surface model of the plasma excitation device. Under high-altitude and high-speed flight conditions, the use of the plasma excitation device can effectively improve the lift-drag characteristics of the aircraft and bring about an increase in the lift coefficient. The plasma excitation device studied in the present invention has the following relationship between the increase in the lift coefficient and the power P and the number N of the plasma excitation device:

[0113]

[0114] Where ΔC L is the lift coefficient increment, in %, N is the number of excitation device groups, P is the excitation power, and under the condition of the presence of plasma excitation device, the aircraft thrust, lift drag and torque coefficient (4) are modified as follows:

[0115]

[0116] in is the distance between the plasma excitation device and the moment reference point of the aircraft, and the equivalent rudder deflection is defined as follows:

[0117]

[0118] In the presence of a plasma excitation device, the longitudinal motion plane model (13) of the hypersonic vehicle is modified to

[0119]

[0120] The definitions of other variables in the above model are the same as those in model (13). Models (16) and (17) show that the function of the plasma excitation device is equivalent to the deflection of the control surface. Under the condition of thin atmosphere at high altitude, the plasma excitation device releases plasma into space with a certain power and number of groups, which can increase the efficiency of the control surface and improve the lift-drag characteristics of the aircraft.

[0121] Step S4, design of nonlinear disturbance observer. Nonlinear disturbance observer (NDO) technology is a new tool that has emerged in recent years to improve the robustness of control systems. NDO is used to smoothly estimate the uncertainties of the model (perturbation parameters, external disturbances, unmodeled dynamics, etc.) and compensate them in the control law. This can not only ensure good robustness of the control law, but also overcome the conservative shortcomings of traditional robust control and the instability problem in controller design. In order to overcome the conservative shortcomings of traditional robust control, the present invention designs a new type of nonlinear disturbance observer. In order to facilitate the design of the nonlinear disturbance observer, the following nonlinear system is first considered.

[0122]

[0123] For the above system, there is the following lemma:

[0124] Lemma 1: For system (18), the following nonlinear system is designed to ensure that the system input estimation error converges to a small neighborhood of zero in a finite time:

[0125]

[0126] Where R,a i is the parameter to be designed, then there exists φ>0,ιφ>n such that

[0127]

[0128] In the formula, O((1R) ιφ-i+1 ) indicates ζ i With i-1 The approximation of (t) is O((1R) ιφ-i+1 ) level, Based on Lemma 1, a new type of nonlinear disturbance observer is designed. Without loss of generality, consider the following uncertain dynamic system

[0129]

[0130] Where v(t)∈R is the system state variable, F(t), G(t) are continuous functions, d(t) is the system uncertainty disturbance, and u(t) is the system input variable; Lemma 2: For the uncertain system (21), the following nonlinear disturbance observer is designed:

[0131]

[0132] in are the estimated values ​​of the system states v(t) and d(t), respectively. If R→+∞, then

[0133]

[0134] Step S5, based on the speed control law design of the disturbance observer, the actual control target of the hypersonic vehicle is: by adjusting the ramjet engine oil-air ratio δ T and the pitch control surface deflection angle Dz, in the longitudinal plane, the speed and altitude are robustly tracked to their respective reference inputs, and the flight attitude is stabilized and elastic vibration is suppressed. According to the aircraft dynamics model, due to the oil-gas ratio δ T It directly affects the thrust, so the speed change is mainly controlled by the oil-gas ratio, and the altitude change is mainly controlled by Dz. This is because Dz directly affects the pitch rate of the aircraft, which in turn affects the pitch angle and track angle, and finally controls the change of the flight altitude. Therefore, in order to facilitate the design of the control law and improve the robustness of the control system, the longitudinal motion model of the aircraft is usually decomposed into a speed subsystem and an altitude subsystem in form. A nonlinear control law that can estimate external disturbances and model uncertainty factors is used. The control law is designed for the speed subsystem based on the dynamic inverse theory to achieve robust tracking of the speed to the reference input. The speed tracking error is defined as follows:

[0135]

[0136] Taking the first-order derivative of the above formula, we can get

[0137]

[0138] The speed control law is designed as

[0139]

[0140] Among them, k v1 ,k v2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d v Estimated value of

[0141]

[0142] Among them, R V ,a v1,a v2 ∈R + are the parameters to be designed, is the estimated value of V;

[0143] Step S6, based on the disturbance observer, the altitude control law is designed. Based on the inverse design theory, two intermediate virtual control laws and the final actual control law are designed for the altitude subsystem, so that γ approaches γ d , and then achieve the maintenance and tracking of altitude. The first step: define the track angle tracking error

[0144]

[0145] Taking the derivative of the above formula, we have

[0146]

[0147] The virtual control law θ is designed as

[0148]

[0149] Among them, k γ1 ,k γ2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d γ Estimated value of

[0150]

[0151] Among them, R γ ,a γ1 ,a γ2 ∈R + are the parameters to be designed, is the estimated value of γ; Step 2: Define the pitch angle tracking error

[0152]

[0153] The first-order derivative of the above formula is

[0154]

[0155] The virtual control law Q is designed as

[0156]

[0157] Among them, k θ1 ,k θ2 is the parameter to be designed; Step 3: Define the pitch angle rate tracking error

[0158]

[0159] The first-order derivative of the above formula is

[0160]

[0161] The actual control law is designed as

[0162]

[0163] Among them, k Q1 ,k Q2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d Q Estimated value of

[0164]

[0165] Among them, R Q ,a Q1 ,a Q2 ∈R + are the parameters to be designed, is the estimated value of Q;

[0166] Step S7, verification of the effectiveness of the control algorithm, see Figure 4-Figure 13 , taking the dynamic model described in (17) as the controlled object, the nonlinear control strategy designed by the present invention is adopted to realize the tracking control of the altitude and speed instructions of the hypersonic aircraft in the cruise phase, wherein the tracking control instructions are designed as constant reference signals, as follows

[0167]

[0168] Among them, H ref0 and V ref0 is the initial flight altitude and speed of the aircraft.

[0169] The maximum engine valve opening is set to 1.5, the maximum rudder deflection is 20°, and the initial state of the aircraft is set to

[0170] H ref0 =30km,V ref0 =2500m / s,γ 0 =0.5°,θ 0 =4°,q 0 =0.001° / s

[0171] The aerodynamic, torque and engine parameters of the aircraft are:

[0172]

[0173] The controller parameters are set to

[0174]

[0175] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device.

[0176] Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and variations may be made to the embodiments without departing from the principles and spirit of the present invention, and that the scope of the present invention is defined by the appended claims and their equivalents.

Claims

1. A cruise tracking control method for a hypersonic vehicle with a plasma excitation device, characterized in that: The steps include: Step S1, establishing a longitudinal motion plane dynamics model of the aircraft; Step S2, control-oriented longitudinal motion plane model conversion; Step S3, establishing an equivalent control surface model of the plasma excitation device; Step S4, design of nonlinear disturbance observer; Step S5, designing a speed control law based on a disturbance observer; Step S6, designing a height control law based on a disturbance observer; Step S7, verification of the effectiveness of the control algorithm; In step S1, based on the external force on the aircraft and the rotation relationship around the center of mass, the longitudinal motion plane model of the hypersonic aircraft is as follows: where m, h, V, α, θ, q and J y They are respectively the mass of the aircraft, flight altitude, speed, flight angle of attack, trajectory inclination, pitch rate and moment of inertia around the Y axis of the missile body, T, L, D, M t They are engine thrust, aerodynamic lift, drag and pitch moment respectively. The specific expressions are: Among them C T ,C L ,C D ,C m are the engine thrust coefficient, lift coefficient, drag coefficient and pitch moment coefficient, and their values ​​are related to the engine valve opening δ′ T , flight attack angle α, pitch control surface deflection angle δ′ z related, S t ,S ref ,z t , They are dynamic pressure, engine thrust reference area, aerodynamic force reference area, distance from engine thrust to vehicle moment reference point, and aerodynamic moment length; The hypersonic vehicle longitudinal motion plane model after the control-oriented longitudinal motion plane model conversion in step S2 is: The lift coefficient increment of the plasma excitation device in step S3 has the following relationship with the excitation power P of the plasma excitation device and the number of excitation device groups N: Where ΔC L is the lift coefficient increment, in %, N is the number of excitation device groups, P is the excitation power, and under the condition of the presence of plasma excitation device, the thrust, lift and torque coefficients of the aircraft (4) are modified as follows: in is the distance between the plasma excitation device and the moment reference point of the aircraft, and the equivalent rudder deflection is defined as follows: In the presence of a plasma excitation device, the hypersonic vehicle longitudinal motion plane model (13) is modified to:

2. A cruise tracking control method for a hypersonic vehicle with a plasma excitation device according to claim 1, characterized in that: The nonlinear disturbance observer design in step S4 considers the following uncertain dynamic system: Where v(t)∈R is the system state variable, F(t), G(t) are continuous functions, d(t) is the system uncertainty disturbance, u(t) is the system input variable. For the uncertain system (21), the designed nonlinear disturbance observer is as follows: in are the estimated values ​​of the system states v(t) and d(t), respectively. If R→+∞, then 3. A cruise tracking control method for a hypersonic vehicle with a plasma excitation device according to claim 2, characterized in that: The speed control law of step S5 is designed as follows: Among them, k v1 ,k v2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d v The estimated value of Among them, R V ,a v1 ,a v2 ∈R + are the parameters to be designed, is the estimated value of V.

4. A cruise tracking control method for a hypersonic vehicle with a plasma excitation device according to claim 3, characterized in that: The altitude control law design in step S6 is to design two intermediate virtual control laws and the final actual control law for the altitude subsystem to achieve altitude maintenance and tracking, wherein one virtual control law θ is designed as: Among them, k γ1 ,k γ2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d γ The estimated value of Among them, R γ ,a γ1 ,a γ2 ∈R + are the parameters to be designed, is the estimated value of γ, and another virtual control law Q is designed as, Among them, k θ1 ,k θ2 are the parameters to be designed, and the actual control law is designed as follows: Among them, k Q1 ,k Q2 is the parameter to be designed, and is the nonlinear disturbance observer NDO for the noise d Q The estimated value of Among them, R Q ,a Q1 ,a Q2 ∈R + are the parameters to be designed, is the estimated value of Q.

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