An Aircraft Dynamic RCS Simulation Method for En Route Flight

Abstract

The present invention discloses an aircraft dynamic RCS simulation method for the en-route flight process, which relates to the technical fields of simulation and radar detection, and includes the following steps: establishing a six-degree-of-freedom dynamics model and control law of the aircraft according to various parameter data of the aircraft to be simulated, and simulating the flight route of the aircraft to obtain flight route simulation data; establishing a calculation model of the azimuth angle and altitude angle of the radar relative to the aircraft based on the line-of-sight angle range of the aircraft's body coordinate system according to the flight route simulation data of the aircraft; calculating the RCS of the aircraft in different flight postures and establishing a dynamic RCS library; using the dynamic RCS library to simulate the six-degree-of-freedom flight track of the aircraft according to the latitude, longitude and altitude of the aircraft's waypoints and the latitude, longitude and altitude of the radar, and calculating the dynamic RCS simulation value of the aircraft. The present invention realizes the analysis of the dynamic RCS of the entire en-route flight process of the aircraft, and provides certain theoretical support for the stealth performance design and test evaluation of the aircraft.

Description

Technical Field

[0001] The present invention relates to the technical fields of simulation technology and radar detection, and particularly to an aircraft dynamic RCS simulation method for the en-route flight process. Background Art

[0002] The radar cross section (RCS) of an aircraft refers to the effective reflection area of the aircraft for radar waves. The dynamic RCS of an aircraft is an important factor affecting the penetration probability of a fighter aircraft. The dynamic RCS simulation of space targets can provide a data basis for target characteristic analysis and recognition, etc. Due to the huge cost of flight tests, it is impossible to carry out research on the radar characteristics of space targets through a large number of experiments. Therefore, simulating radar echo data through simulation calculations to train and test the detection, tracking, and recognition capabilities of radars has become an essential task. The simulation is divided into two steps, one is motion modeling simulation, and the other is electromagnetic scattering characteristic simulation.

[0003] When existing dynamic RCS modeling and simulation methods perform motion modeling and simulation, most are based on a three-degree-of-freedom dynamics model, and generally, the en-route flight process is decomposed into several typical flight segments (such as side-station level flight, back-station pull-up, etc.) for separate modeling. On the one hand, accurate aircraft attitude angles cannot be obtained in the three-degree-of-freedom dynamics model, and the attitude angles can only be approximately estimated based on the track angle. On the other hand, due to the lack of modeling of the entire en-route flight process in existing dynamic RCS modeling, the dynamic RCS of the entire en-route flight process cannot be analyzed during actual use, and the existing results do not comprehensively consider the influence of aircraft attitude angle information on RCS. Only the line-of-sight pitch angle and deflection angle are used as variables to calculate RCS, which is obviously insufficient and reduces the accuracy and flexibility of RCS calculation. Summary of the Invention

[0004] The purpose of the present invention is to provide an aircraft dynamic RCS simulation method for the en-route flight process, which can analyze the dynamic RCS of the entire en-route flight process of an aircraft and provide certain theoretical support for the stealth performance design and test evaluation of the aircraft.

[0005] The present invention provides an aircraft dynamic RCS simulation method for the en-route flight process, including the following steps:

[0006] According to various parameter data of the aircraft to be simulated, establish a six-degree-of-freedom dynamics model and control law of the aircraft, and simulate the flight route of the aircraft through the six-degree-of-freedom dynamics model and control law to obtain flight route simulation data;

[0007] According to the flight route simulation data of the aircraft, establish a calculation model for the azimuth angle and altitude angle of the radar relative to the aircraft based on the line-of-sight angle range of the aircraft's body coordinate system;

[0008] Calculate the RCS of an aircraft in different flight postures using the calculation models of the azimuth angle and altitude angle of the radar relative to the aircraft, and establish a dynamic RCS library;

[0009] According to the latitude, longitude, and altitude of the aircraft's waypoints and the latitude, longitude, and altitude of the radar, use the dynamic RCS library to simulate the six-degree-of-freedom trajectory of the aircraft and calculate the simulated value of the dynamic RCS of the aircraft.

[0010] Furthermore, the various parameter data of the aircraft to be simulated include: the external shape structure of the aircraft and the material type of the aircraft;

[0011] The establishment of the six-degree-of-freedom dynamics model of the aircraft includes the following steps:

[0012] Establish a model based on the external shape structure and material type of the aircraft, and calculate the inertial parameters of the aircraft and the aerodynamic parameters of the aircraft in different flight states according to this model;

[0013] Establish a six-degree-of-freedom dynamics model of the aircraft based on the inertial parameters of the aircraft and the aerodynamic parameters of the aircraft in different flight states.

[0014] Furthermore, the design of the control law of the aircraft using the multi-loop design method includes the following steps:

[0015] Determine the position loop of the aircraft and design the control law of the position loop of the aircraft;

[0016] Determine the trajectory loop of the aircraft and design the control law of the trajectory loop of the aircraft using the PID control method;

[0017] Determine the attitude loop of the aircraft and design the control law of the attitude loop of the aircraft using the method of dynamic inversion;

[0018] Determine the angular velocity loop of the aircraft and design the control law of the angular velocity loop of the aircraft using the method of incremental dynamic inversion.

[0019] Furthermore, the establishment of the calculation models of the azimuth angle and altitude angle of the radar relative to the aircraft based on the line-of-sight angle range of the aircraft's body coordinate system according to the flight route simulation data of the aircraft includes:

[0020] Use the six-degree-of-freedom dynamics model of the aircraft to obtain the attitude angle and position information during the flight of the aircraft;

[0021] Calculate the azimuth angle and altitude angle of the radar relative to the aircraft during the flight of the aircraft according to the attitude angle and position information during the flight of the aircraft, and establish the calculation models of the azimuth angle and altitude angle of the radar relative to the aircraft.

[0022] Furthermore, the position loop of the aircraft is:

[0023]

[0024] Where L is the longitude of the aircraft; B is the latitude of the aircraft; H is the altitude of the aircraft;

[0025] V is the speed of the aircraft; θ is the pitch angle of the aircraft; χ is the azimuth angle of the aircraft;

[0026] The control law of the position loop of the designed aircraft includes:

[0027] Convert the longitude, latitude, and altitude (L0, B0, H0) of the radar into coordinates (X0, Y0, Z0) in the geocentric rectangular coordinate system, then:

[0028]

[0029] Where L0 is the longitude of the radar; B0 is the latitude of the radar; H0 is the altitude of the radar;

[0030] According to the formula:

[0031]

[0032] The longitude and latitude (L, B, H) of the current point of the aircraft can be converted into coordinates (x, y, z) in the geocentric rectangular coordinate system, and the longitude and latitude (L t , B t , H t ) of the target point of the aircraft can be converted into coordinates (x t , y t , z t ) in the geocentric rectangular coordinate system;

[0033] Where L t is the longitude of the target point of the aircraft; B t is the latitude of the target point of the aircraft; H t is the altitude of the target point of the aircraft;

[0034] Let:

[0035]

[0036] Then the control law of the position loop of the aircraft is as follows:

[0037]

[0038] Where χ and γ are the azimuth angle and the track angle of the aircraft respectively; χ γ is the reference signal of the azimuth angle of the aircraft.

[0039] Furthermore, the track loop of the aircraft is:

[0040]

[0041] where ω H / I is the angular velocity of the track system relative to the inertial system, and is:

[0042]

[0043] Substituting (7) into (6), we get:

[0044]

[0045] where m is the mass of the aircraft;

[0046] are the components of the resultant external force of the aircraft on each coordinate axis in the track system, and:

[0047]

[0048] where C H / B is the transformation matrix for coordinate transformation from the local system to the track system, and has:

[0049]

[0050] where β is the sideslip angle of the aircraft; using s to represent sin and c to represent cos;

[0051] are the components of the resultant external force of the aircraft on each coordinate axis in the local system. The resultant external force of the aircraft includes aerodynamic force and gravity, and has:

[0052]

[0053] where f Ax , f Ay , f Az are the components of the aerodynamic force on each coordinate axis in the local system, and their expressions are respectively:

[0054]

[0055]

[0056]

[0057] where S ref is the wing reference area; is the dynamic pressure;

[0058] C X0 , C xα , C Xq , C Yβ , C Yγ , C Yp , C Z0 , C Zα, C Zq are all aircraft aerodynamic derivatives;

[0059] where C B / I is the coordinate transformation matrix from the inertial system to the local system, and we have:

[0060]

[0061] where C H / I is the transformation matrix from the inertial system to the flight path coordinate system;

[0062] The flight path law for designing the position loop of the aircraft using the PID control method includes:

[0063] Using PID control to control the engine thrust T of the aircraft, we have:

[0064] T = K vp (V r - V) + K vI I ev

[0065]

[0066] where V r is the reference speed; K vp , K vI are the controller gains;

[0067] is the difference between the reference speed and the aircraft speed;

[0068] Set the turning mode of the aircraft to be achieved by the roll of the aircraft, and use PID control to control the turning of the aircraft. Then the reference signal μ des of the aircraft bank angle μ is as follows:

[0069]

[0070]

[0071] where μ is the aircraft bank angle; μ des is the reference signal of the aircraft bank angle μ;

[0072] χ r is the reference angle of the aircraft azimuth angle χ given by the azimuth loop;

[0073] K μp , K μI are the controller gains;

[0074] is the difference between the reference angle of the aircraft azimuth angle χ and the aircraft azimuth angle χ;

[0075] where Rate and amplitude saturation function set for the bank angle μ of the aircraft:

[0076]

[0077] where,

[0078]

[0079]

[0080] where μ correspond to the maximum and minimum allowable ranges of the bank angle μ of the aircraft respectively;

[0081] μ max and μ min are the maximum and minimum values of the allowable bank angle μ of the aircraft respectively;

[0082] μ0 is the μ of the previous control period of the controller; d ct is the controller period;

[0083] is the maximum rate limit of μ;

[0084] The incremental dynamic inversion method is used to design α des as follows:

[0085]

[0086]

[0087]

[0088] where α is the angle of attack of the aircraft; α des is the reference signal of the angle of attack α of the aircraft;

[0089] γ γ is the reference angle of the flight path angle γ of the aircraft obtained by the flight path loop;

[0090] K αp 、K αI 、K αp 、K αI are the controller gains respectively;

[0091] is the difference between the reference angle of the flight path angle γ of the aircraft and the flight path angle γ of the aircraft;

[0092] α0 is the angle of attack of the aircraft in the previous control period; Δα is the calculated control increment; γ0 is the flight path angle of the aircraft in the previous control period;

[0093] where,

[0094] α0(t) = α(t - d ct )

[0095] γ0(t) = γ(t - d ct ) (20)

[0096] where d ct is the control period; t is the current time step;

[0097] where is a rate and amplitude saturation function set for the angle of attack α of the aircraft:

[0098]

[0099] where α are the maximum and minimum allowable ranges of the angle of attack α, respectively, and they are respectively:

[0100]

[0101]

[0102] where α max and α min are the maximum and minimum amplitude limits of the angle of attack α of the aircraft, respectively;

[0103] is the maximum rate limit of the angle of attack α of the aircraft.

[0104] Furthermore, the attitude loop of the aircraft is:

[0105]

[0106] where is the angular velocity vector of the aircraft's body frame relative to the inertial frame, and its expression is:

[0107]

[0108] where p, q, and r are the angular rates of rotation about the three axes of the aircraft's body frame, respectively;

[0109] where is the angular velocity vector of the aircraft's body frame relative to the trajectory coordinate system, and its expression is:

[0110]

[0111] where is the angular velocity vector of the trajectory coordinate system relative to the inertial coordinate system, and its expression is:

[0112]

[0113] where C B / H is the transformation matrix for converting the track coordinate system to this system, and its expression is:

[0114]

[0115] C H / B is the transformation matrix for converting the coordinates in this system to the track system;

[0116] The control law of the attitude loop of the aircraft is designed by using the method of dynamic inversion, including:

[0117]

[0118]

[0119]

[0120]

[0121] where p des , q des , r des are respectively the reference signals of the angular rates p, q, r of rotation about the three axes of the aircraft's own system; K μp , K αp , K βp , K μI , K αI , K βI are respectively the controller gains;

[0122] I eμ , I eα , I eβ , I eλ are the accumulated errors in the integrator;

[0123] μ r is the reference signal of the bank angle μ of the aircraft; α r is the reference signal of the angle of attack α of the aircraft;

[0124] β r is the reference signal of the sideslip angle β of the aircraft.

[0125] Furthermore, the angular velocity loop of the aircraft is:

[0126]

[0127] L, m, n are respectively the torques about the x, y, z axes of this system, and their expressions are respectively:

[0128] l = l a + l G + l I(31)

[0129] m = m a + m G + m I (32)

[0130] n = n a + n G + n I (33)

[0131] where l a , m a , n a are the aerodynamic moments of the aerodynamic force on each coordinate axis of this system, and their expressions are respectively:

[0132]

[0133]

[0134]

[0135] where C lβ , C lp , C lr , C m0 , C mα , C mq , C nβ , C np , C nr , are all aircraft aerodynamic derivatives;

[0136] δ α is the deflection angle of the aircraft aileron;

[0137] δ r is the deflection angle of the aircraft rudder; δ e is the deflection angle of the aircraft elevator; is the aerodynamic chord length of the aircraft;

[0138] where l G , m G , n G are the aerodynamic moments of the three axes caused by gravity, and their expressions are respectively:

[0139]

[0140] where S x is the x-axis coordinate of the center of mass in this system;

[0141] S y is the y-axis coordinate of the center of mass in this system;

[0142] S z is the z-axis coordinate of the centroid in this system;

[0143] where l I , m I , n I are the three-axis components of the moment of inertia in this system, and their expressions are respectively:

[0144]

[0145] where S = [S x , S y , S z T ; ω = [p, q, r] T ; v is the aircraft velocity vector;

[0146] ρ i is the position vector of the aircraft mass element relative to the origin of the aircraft's own system;

[0147] The control law of the aircraft's angular velocity loop is designed by using the incremental dynamic inversion method, including:

[0148]

[0149] where J is the aircraft moment of inertia;

[0150]

[0151]

[0152] where p0, q0, r0 are the values of p, q, r in the previous control period, and there are

[0153] p0(t) = p(t - d ct )

[0154] q0(t) = q(t - d ct )

[0155] r0(t) = r(t - d ct ) (39)

[0156] where,

[0157]

[0158] where δ a is the aileron deflection angle; δ e is the elevator deflection angle; δ γ is the rudder deflection angle;

[0159] Δδ a is the aileron control increment; Δδ e ​is the elevator control increment; Δδ γ is the rudder control increment;

[0160] δ α0 is the value of the aileron deflection angle in the previous control cycle; δ e0 is the value of the elevator deflection angle in the previous control cycle; δ γ0 is the value of the rudder deflection angle in the previous control cycle, and respectively there are:

[0161]

[0162] where t is the current time step, d ct is the control cycle.

[0163] Furthermore, the method for calculating the RCS of the aircraft in different flight attitudes by using the calculation model of the azimuth angle and altitude angle of the radar relative to the aircraft, and establishing a dynamic RCS library includes the following steps:

[0164] Convert the waypoints of the aircraft and the coordinates of the radar to the coordinates in the Earth station centered coordinate system, and perform flight route simulation of the aircraft according to the coordinates in the Earth station centered coordinate system to generate flight data of the aircraft;

[0165] Calculate the azimuth angle and altitude angle of the radar in the body coordinate system of the aircraft through coordinate transformation according to the Euler angle and position information of the aircraft in the flight data of the aircraft. The calculation formulas for the azimuth angle and altitude angle of the radar are:

[0166]

[0167] where is the azimuth angle of the radar relative to the aircraft, and its value range is [-π, π];

[0168] θ′ is the altitude angle of the radar relative to the aircraft, and its value range is [0, π];

[0169] where (x4, y4, z4) is the coordinate of the radar in the body coordinate system, and is:

[0170]

[0171] where (x T , y T , z T ) is the coordinate of the radar in the Earth station centered coordinate system;

[0172] Q is the transformation matrix from the radar coordinate system to the body coordinate system, and is:

[0173]

[0174] where θ is the pitch angle, ψ is the yaw angle, and φ is the roll angle, all of which are the attitude angles of the aircraft;

[0175] where (x p , y p , z p ) are the coordinates of the trajectory P of the aircraft;

[0176] Calculate the dynamic RCS value of the aircraft at each moment according to the azimuth angle and altitude angle information of the radar in the body coordinate system of the aircraft, and obtain the dynamic RCS library.

[0177] Compared with the prior art, the beneficial effects of the present invention are:

[0178] The method for modeling the flight dynamic RCS based on the six-degree-of-freedom route simulation model proposed by the present invention improves the problem that the calculation of the azimuth angle and altitude angle in the existing literature can only be carried out in a limited angular domain by using the method of establishing a dynamic RCS library based on the plane wave method, and gives clear and accurate definitions of the azimuth angle and altitude angle in the full angular domain. The results calculated based on the azimuth angle and altitude angle are sorted into a two-dimensional RCS data table of RCS with respect to the altitude angle and azimuth angle to obtain the dynamic RCS library, realizing the analysis of the dynamic RCS of the entire route flight process of the aircraft, and providing a certain theoretical support for the stealth performance design and test evaluation of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0179] The drawings are used to provide a further understanding of the present invention and constitute a part of the specification. Together with the embodiments of the present invention, they are used to explain the present invention and do not constitute a limitation to the present invention. In the drawings:

[0180] Figure 1 is the principle flowchart of a method for simulating the dynamic RCS of an aircraft for a route flight process proposed by the present invention;

[0181] Figure 2 is the transformation relationship diagram between various coordinate systems in a method for simulating the dynamic RCS of an aircraft for a route flight process proposed by the present invention;

[0182] Figure 3 is the schematic diagram of the topocentric rectangular coordinate system of a method for simulating the dynamic RCS of an aircraft for a route flight process proposed by the present invention;

[0183] Figure 4 is the schematic diagram of the geocentric rectangular coordinate system of a method for simulating the dynamic RCS of an aircraft for a route flight process proposed by the present invention;

[0184] Figure 5 is the schematic diagram of the altitude angle and azimuth angle in a method for simulating the dynamic RCS of an aircraft for a route flight process proposed by the present invention;

[0185] Figure 6 It is the calculation flowchart of the dynamic RCS in the six-degree-of-freedom trajectory simulation in an aircraft dynamic RCS simulation method for the en-route flight process proposed by the present invention;

[0186] Figure 7 It is the graph showing the change of aircraft attitude simulation over time in an embodiment of an aircraft dynamic RCS simulation method for the en-route flight process proposed by the present invention;

[0187] Figure 8 It is the graph showing the relationship between the aircraft trajectory and the position of the aircraft and the radar in an embodiment of an aircraft dynamic RCS simulation method for the en-route flight process proposed by the present invention;

[0188] Figure 9 It is the dynamic RCS graph in the aircraft trajectory simulation of an aircraft dynamic RCS simulation method for the en-route flight process proposed by the present invention. Detailed implementation manners

[0189] Next, in combination with the accompanying drawings in the embodiments of the present invention, the technical solutions in the embodiments of the present invention will be clearly and completely described. However, it should be understood that the protection scope of the present invention is not limited by the specific implementation manners.

[0190] Embodiment

[0191] As Figure 1-8 shown, an aircraft dynamic RCS simulation method for the en-route flight process includes the following steps:

[0192] Step 1: According to the parameter data of the aircraft to be simulated, establish a six-degree-of-freedom dynamics model and control law of the aircraft, and simulate the flight route of the aircraft through the six-degree-of-freedom dynamics model and control law to obtain flight route simulation data.

[0193] The parameter data of the aircraft to be simulated includes: the external shape structure of the aircraft, the material type of the aircraft.

[0194] Step 1.1: Establish a six-degree-of-freedom dynamics model of the aircraft, including the following steps:

[0195] Establish a model according to the external shape structure and material type of the aircraft, and calculate the inertial parameters of the aircraft and the aerodynamic parameters of the aircraft in different flight states based on this model. Specifically, establish a model of the external shape structure and material of the aircraft based on CAD, and import this model into Solidworks software to calculate the inertial parameters of the aircraft, and use CFD to calculate the aerodynamic parameters of the aircraft in different flight states.

[0196] Establish a six-degree-of-freedom dynamic model of the aircraft based on the inertial parameters of the aircraft and the aerodynamic parameters of the aircraft in different flight states. The aerodynamic parameters are all obtained by interpolation of the current flight state.

[0197] Step 1.2: Design the control law of the aircraft using the multi-loop design method, including the following steps:

[0198] For the convenience of control law design, the North Celestial East coordinate system is adopted, and the aircraft model is converted into a loop cascade form, where the geometric relationships between the coordinate systems are as Figure 2 shown.

[0199] Step 1.2.1: Determine the position loop of the aircraft and design the control law of the position loop of the aircraft, specifically as follows:

[0200] The position loop of the aircraft is:

[0201]

[0202] where L is the longitude of the aircraft; B is the latitude of the aircraft; H is the altitude of the aircraft;

[0203] V is the speed of the aircraft; θ is the pitch angle of the aircraft; χ is the azimuth angle of the aircraft;

[0204] The control law for designing the position loop of the aircraft includes:

[0205] Convert the longitude, latitude, and altitude (L0, B0, H0) of the radar to the coordinates (X0, Y0, Z0) in the geocentric rectangular coordinate system, then:

[0206]

[0207] where L0 is the longitude of the radar; B0 is the latitude of the radar; H0 is the altitude of the radar;

[0208] According to the formula:

[0209]

[0210] The longitude, latitude, and altitude (L, B, H) of the current point of the aircraft can be converted to the coordinates (x, y, z) in the geocentric rectangular coordinate system, and the longitude, latitude, and altitude (L t , B t , H t ) of the target point of the aircraft can be converted to the coordinates (x t , y t , z t ) in the geocentric rectangular coordinate system;

[0211] where L t is the longitude of the target point of the aircraft; B t is the latitude of the target point of the aircraft; H tis the altitude of the target point of the aircraft;

[0212] Let:

[0213]

[0214] Then the control law of the position loop of the aircraft is as follows:

[0215]

[0216] where χ and γ are the azimuth angle and the track angle of the aircraft respectively; χ γ is the reference signal of the azimuth angle of the aircraft

[0217] Step 1.2.2: Determine the track loop of the aircraft, and design the control law of the track loop of the aircraft by using the PID control method, specifically as follows:

[0218] The track loop of the aircraft is:

[0219]

[0220] where ω H / I is the angular velocity of the track system relative to the inertial system, and is:

[0221]

[0222] Substitute (7) into (6), and we can get:

[0223]

[0224] where m is the mass of the aircraft;

[0225] are the components of the resultant external force of the aircraft on each coordinate axis of the track system respectively, and:

[0226]

[0227] where C H / B is the transformation matrix for converting the coordinates in this system to the track system, and there is:

[0228]

[0229] where β is the sideslip angle of the aircraft; use s to represent sin and c to represent cos;

[0230] are the components of the resultant external force of the aircraft on each coordinate axis of this system respectively. The resultant external force of the aircraft includes aerodynamic force and gravity, and there is:

[0231]

[0232] where fAx , f Ay , f Az are the components of the aerodynamic force on each coordinate axis of this system, and their expressions are respectively:

[0233]

[0234]

[0235]

[0236] where S ref is the wing reference area; is the dynamic pressure;

[0237] C X0 , C Xα , C Xq , C Yβ , C Yγ , C Yp , C Z0 , C Zα , C Zq are all aircraft aerodynamic derivatives;

[0238] where C B / I is the coordinate transformation matrix from the inertial system to this system, and there is:

[0239]

[0240] where C H / I is the transformation matrix from the inertial system to the flight path coordinate system;

[0241] The flight path law for designing the position loop of the aircraft using the PID control method includes:

[0242] Using PID control to control the engine thrust T of the aircraft, there is:

[0243] T = K vp (V r - V) + K vI I ev

[0244]

[0245] where V r is the reference speed; K vp , K vI are the controller gains;

[0246] is the difference between the reference speed and the aircraft speed;

[0247] Set the steering method of the aircraft to be achieved through the roll of the aircraft, and use PID control to control the steering of the aircraft. Then the reference signal μ of the aircraft tilt angle μ des is as follows:

[0248]

[0249]

[0250] where μ is the tilt angle of the aircraft; μ des is the reference signal of the aircraft tilt angle μ;

[0251] χ r is the reference angle of the aircraft azimuth angle χ given by the azimuth loop;

[0252] K μp 、K μI are controller gains;

[0253] is the difference between the reference angle of the aircraft azimuth angle χ and the aircraft azimuth angle χ;

[0254] where is the rate and amplitude saturation function set for the tilt angle μ of the aircraft:

[0255]

[0256] where,

[0257]

[0258]

[0259] where μ corresponds to the maximum and minimum allowable ranges of the aircraft tilt angle μ respectively;

[0260] μ max 、μ min are the maximum and minimum values of the allowable aircraft tilt angle μ respectively;

[0261] μ0 is μ in the previous control cycle of the controller; d ct is the controller period;

[0262] is the maximum rate limit of μ;

[0263] Use the method of incremental dynamic inversion to design α des as follows:

[0264]

[0265]

[0266]

[0267] where α is the angle of attack of the aircraft; α des is the reference signal of the angle of attack α of the aircraft;

[0268] γ γ is the reference angle of the flight path angle γ of the aircraft obtained by the flight path loop;

[0269] K αp 、K αI 、K αp 、K αI are the controller gains respectively;

[0270] is the difference between the reference angle of the flight path angle γ of the aircraft and the flight path angle γ of the aircraft;

[0271] α0 is the angle of attack of the aircraft in the previous control period; Δα is the calculated control increment; γ0 is the flight path angle of the aircraft in the previous control period;

[0272] where,

[0273] α0(t) = α(t - d ct )

[0274] γ0(t) = γ(t - d ct ) (20)

[0275] where d ct is the control period; t is the current time step;

[0276] where is the rate and amplitude saturation function set for the angle of attack α of the aircraft:

[0277]

[0278] where α are the maximum and minimum allowable ranges of the angle of attack α respectively, which are:

[0279]

[0280]

[0281] where α max 、α min are the maximum and minimum amplitude limits of the angle of attack α of the aircraft respectively;

[0282] is the maximum rate limit of the angle of attack α of the aircraft.

[0283] Step 1.2.3: Determine the attitude loop of the aircraft and design the control law of the attitude loop of the aircraft by using the dynamic inversion method, specifically as follows:

[0284] The attitude loop of the aircraft is:

[0285]

[0286] where is the angular velocity vector of the aircraft's body frame relative to the inertial frame, and its expression is:

[0287]

[0288] where p, q, and r are the angular rates of rotation about the three axes of the aircraft's body frame respectively;

[0289] where is the angular velocity vector of the aircraft's body frame relative to the trajectory frame, and its expression is:

[0290]

[0291] where is the angular velocity vector of the trajectory frame relative to the inertial frame, and its expression is:

[0292]

[0293] where C B / H is the transformation matrix from the trajectory frame to the body frame, and its expression is:

[0294]

[0295] C H / B is the transformation matrix for coordinate transformation from the body frame to the trajectory frame;

[0296] Design the control law of the attitude loop of the aircraft by using the dynamic inversion method, including:

[0297]

[0298]

[0299]

[0300]

[0301] where p des 、q des 、r des are the reference signals of the angular rates p, q, and r of rotation about the three axes of the aircraft's body frame respectively; K μp ,K αp ,Kβp , K μI , K αI , K βI are the controller gains respectively;

[0302] I eμ , I eα , I eβ , I eλ is the cumulative error in the integrator;

[0303] μ r is the reference signal for the bank angle μ of the aircraft; α r is the reference signal for the angle of attack α of the aircraft;

[0304] β r is the reference signal for the sideslip angle β of the aircraft.

[0305] Step 1.2.4: Determine the angular velocity loop of the aircraft, and design the control law of the angular velocity loop of the aircraft by using the incremental dynamic inversion method, specifically as follows:

[0306] The angular velocity loop of the aircraft is:

[0307]

[0308] L, m, and n are the torques about the x, y, and z axes of this system respectively, and their expressions are respectively:

[0309] l = l a + l G + l I (31)

[0310] m = m a + m G + m I (32)

[0311] n = n a + n G + n I (33)

[0312] where l a , m a , n a are the aerodynamic torques on each coordinate axis of this system, and their expressions are respectively:

[0313]

[0314]

[0315]

[0316] where C lβ , Clp , C lr , C m0 , C mα , C mq , C nβ , C np , C mr , are all aircraft aerodynamic derivatives;

[0317] δ α is the deflection angle of the aircraft aileron;

[0318] δ r is the deflection angle of the aircraft rudder; δ e is the deflection angle of the aircraft elevator; is the aerodynamic chord length of the aircraft;

[0319] where l G , m G , n G are the aerodynamic moments of the three axes caused by gravity, and their expressions are respectively:

[0320]

[0321] where S x is the coordinate of the center of mass on the x-axis of this system;

[0322] S y is the coordinate of the center of mass on the y-axis of this system;

[0323] S z is the coordinate of the center of mass on the z-axis of this system;

[0324] where l I , m I , n I are the components of the inertia moment in the three axes of this system, and their expressions are respectively:

[0325]

[0326] where S = [S x , S y , S z T ; ω = [p, q, r] T ; v is the aircraft velocity vector;

[0327] ρ i is the position vector of the aircraft mass element relative to the origin of the aircraft's own system;

[0328] The control law of the aircraft angular velocity loop is designed by using the incremental dynamic inversion method, including: ​

[0329]

[0330] where J is the moment of inertia of the aircraft;

[0331]

[0332]

[0333] where p0, q0, r0 are the values of p, q, r in the previous control period, and there are

[0334] p0(t) = p(t - d ct )

[0335] q0(t) = q(t - d ct )

[0336] r0(t) = r(t - d ct ) (39)

[0337] where,

[0338]

[0339] where δ a is the aileron deflection angle; δ e is the elevator deflection angle; δ γ is the rudder deflection angle;

[0340] Δδ a is the aileron control increment; Δδ e is the elevator control increment; Δδ γ is the rudder control increment;

[0341] δ α0 is the value of the aileron deflection angle in the previous control period; δ e0 is the value of the elevator deflection angle in the previous control period; δ γ0 is the value of the rudder deflection angle in the previous control period, and respectively there are:

[0342]

[0343] where t is the current time step, d ct is the control period.

[0344] Step 2: According to the flight route simulation data of the aircraft, establish a calculation model for the azimuth and altitude angles of the radar relative to the aircraft based on the line-of-sight angle range of the aircraft's body coordinate system, including:

[0345] Use the six-degree-of-freedom dynamic model of the aircraft to obtain the attitude angle and position information during the flight of the aircraft;

[0346] Calculate the azimuth angle and elevation angle of the radar relative to the aircraft during flight based on the attitude angle and position information of the aircraft during flight, and establish a calculation model for the azimuth angle and elevation angle of the radar relative to the aircraft.

[0347] Step 3: Use the calculation model of the azimuth angle and elevation angle of the radar relative to the aircraft to calculate the RCS of the aircraft in different flight postures, and establish a dynamic RCS library, as Figure 6 shown, including the following steps:

[0348] Convert the waypoints of the aircraft and the coordinates of the radar to the coordinates in the earth station coordinate system, and perform flight route simulation of the aircraft based on the coordinates in the earth station coordinate system to generate flight data of the aircraft;

[0349] Calculate the azimuth angle and elevation angle of the radar in the body coordinate system of the aircraft through coordinate transformation according to the Euler angle and position information of the aircraft in the flight data of the aircraft. The calculation formulas for the azimuth angle and elevation angle of the radar are:

[0350]

[0351] where is the azimuth angle of the radar relative to the aircraft, and its value range is [-π, π];

[0352] θ′ is the elevation angle of the radar relative to the aircraft, and its value range is [0, π];

[0353] where (x4, y4, z4) are the coordinates of the radar in the body coordinate system, which are:

[0354]

[0355] where (x T , y T , z T ) are the coordinates of the radar in the earth station coordinate system;

[0356] Q is the transformation matrix from the radar coordinate system to the body coordinate system, which is:

[0357]

[0358] where θ is the pitch angle, ψ is the yaw angle, and φ is the roll angle, all of which are the attitude angles of the aircraft;

[0359] where (x p , y p , z p ) are the coordinates of the trajectory P of the aircraft;

[0360] Calculate the dynamic RCS value of the aircraft at each moment according to the azimuth angle and elevation angle information of the radar in the body coordinate system of the aircraft to obtain a dynamic RCS library.

[0361] Step 4: According to the latitude, longitude and altitude of the waypoints of the aircraft and the latitude, longitude and altitude of the radar, use the dynamic RCS library to simulate the six-degree-of-freedom trajectory of the aircraft, and calculate the dynamic RCS simulation value of the aircraft.

[0362] Specifically, the latitude, longitude and altitude of the waypoints of the aircraft and the latitude, longitude and altitude of the radar are used as input data to realize the six-degree-of-freedom trajectory simulation of the aircraft and the real-time simulation calculation of the dynamic RCS of the route.

[0363] In the present invention, a dynamic RCS library is established by using the plane wave method, which improves the problem that the calculation of the azimuth angle and the elevation angle in the existing literature can only be carried out in a limited angular domain, and gives clear and accurate definitions of the azimuth angle and the elevation angle in the full angular domain. The results calculated based on the azimuth angle and the elevation angle are sorted into a two-dimensional RCS data table of RCS versus elevation angle and azimuth angle to obtain the dynamic RCS library.

[0364] The advantage of the present invention is that the latitude, longitude and altitude of the waypoints of the aircraft and the latitude, longitude and altitude of the radar can be directly used as input data, and there is no need for segmented modeling, which is convenient for application in engineering.

[0365] The technical solution of the present invention will be further described below in conjunction with specific embodiments.

[0366] In this embodiment, the aircraft starts from 50°.0′, 0″ east longitude and 60°.0′, 0″ north latitude and passes through successively:

[0367] Waypoint 1: 50°.0′, 0″ east longitude and 60°.12′, 0″ north latitude;

[0368] Waypoint 2: 50°.18′, 0″ east longitude and 60°.12′, 0″ north latitude;

[0369] The flight altitude of the aircraft is 5000 meters, and the speed of the aircraft remains 200 meters per second.

[0370] The position of the radar is at 50°.6′, 0″ east longitude and 60°.6′, 0″ north latitude, and the altitude is 0 meters.

[0371] The simulation process of the aircraft attitude is as Figure 7 shown, the relative relationship between the aircraft trajectory and the radar is as Figure 8 shown, and the dynamic RCS value of the aircraft is as Figure 9 shown.

[0372] Finally, it should be noted that: the above disclosure is only a specific embodiment of the present invention, however, the embodiments of the present invention are not limited thereto, and any changes that can be thought of by those skilled in the art should fall within the protection scope of the present invention.

Claims

1. An aircraft dynamic RCS simulation method for the en-route flight process, characterized in that, It includes the following steps: Based on the various parameter data of the aircraft to be simulated, establish a six-degree-of-freedom dynamics model and control law of the aircraft, and simulate the flight route of the aircraft through the six-degree-of-freedom dynamics model and control law to obtain flight route simulation data; Based on the flight route simulation data of the aircraft, establish a calculation model for the azimuth and elevation angles of the radar relative to the aircraft within the line-of-sight angle range of the aircraft's body coordinate system; Use the calculation model for the azimuth and elevation angles of the radar relative to the aircraft to calculate the RCS of the aircraft in different flight postures, and establish a dynamic RCS library; Based on the latitude, longitude, and altitude of the aircraft's waypoints and the latitude, longitude, and altitude of the radar, use the dynamic RCS library to simulate the six-degree-of-freedom trajectory of the aircraft and calculate the dynamic RCS simulation value of the aircraft.

2. A dynamic RCS simulation method for an aircraft during en-route flight according to claim 1, characterized in that: The various parameter data of the aircraft to be simulated include: the external shape structure of the aircraft and the material type of the aircraft; The establishment of the six-degree-of-freedom dynamics model of the aircraft includes the following steps: Establish a model based on the external shape structure and material type of the aircraft, and calculate the inertial parameters of the aircraft and the aerodynamic parameters of the aircraft in different flight states according to this model; Based on the inertial parameters of the aircraft and the aerodynamic parameters of the aircraft in different flight states, establish a six-degree-of-freedom dynamics model of the aircraft.

3. A method for simulating the dynamic RCS of an aircraft during en-route flight according to claim 2, characterized in that: Use the multi-loop design method to design the control law of the aircraft, including the following steps: Determine the position loop of the aircraft and design the control law of the position loop of the aircraft; Determine the trajectory loop of the aircraft and use the PID control method to design the control law of the trajectory loop of the aircraft; Determine the attitude loop of the aircraft and use the dynamic inversion method to design the control law of the attitude loop of the aircraft; Determine the angular velocity loop of the aircraft and use the incremental dynamic inversion method to design the control law of the angular velocity loop of the aircraft.

4. A method for aircraft dynamic RCS simulation during en-route flight according to claim 3, characterized in that: The establishment of the calculation model for the azimuth and elevation angles of the radar relative to the aircraft within the line-of-sight angle range of the aircraft's body coordinate system based on the flight route simulation data of the aircraft includes: Use the six-degree-of-freedom dynamics model of the aircraft to obtain the attitude angle and position information of the aircraft during flight; According to the attitude angle and position information of the aircraft during flight, calculate the azimuth and elevation angles of the radar relative to the aircraft during flight, and establish a calculation model for the azimuth and elevation angles of the radar relative to the aircraft.

5. A method for aircraft dynamic RCS simulation during en-route flight according to claim 3, characterized in that: The position loop of the aircraft is: where L is the longitude of the aircraft; B is the latitude of the aircraft; H is the altitude of the aircraft; V is the speed of the aircraft; θ is the pitch angle of the aircraft; χ is the azimuth angle of the aircraft; The design of the control law of the position loop of the aircraft includes: Convert the longitude, latitude, and altitude (L0, B0, H0) of the radar into coordinates (X0, Y0, Z0) in the geocentric rectangular coordinate system, then: where L0 is the longitude of the radar; B0 is the latitude of the radar; H0 is the altitude of the radar; According to the formula: The longitude and latitude (L, B, H) of the current point of the aircraft can be converted into coordinates (x, y, z) in the geocentric rectangular coordinate system, and the longitude and latitude (L t , B t , H t ) of the target point of the aircraft can be converted into coordinates (x t , y t , z t ) in the geocentric rectangular coordinate system; where L t is the longitude of the target point of the aircraft; B t is the latitude of the target point of the aircraft; H t is the altitude of the target point of the aircraft; Let: Then the control law of the position loop of the aircraft is as follows: where χ and γ are the azimuth angle and the track angle of the aircraft respectively; χ γ is the reference signal of the azimuth angle of the aircraft.

6. A dynamic RCS simulation method for an aircraft during en-route flight according to claim 5, characterized in that: The trajectory loop of the aircraft is: where ω H / I is the angular velocity of the track system relative to the inertial system, and is: Substitute (7) into (6), and we can get: where m is the mass of the aircraft; They are the components of the resultant external force of the aircraft on each coordinate axis of the flight path system, and: Among which C H / B is the transformation matrix for coordinate conversion in this system to the track system, and there is: where β is the sideslip angle of the aircraft; use s to represent sin and c to represent cos; They are the components of the resultant external force of the aircraft on each coordinate axis of this system. The resultant external force of the aircraft includes aerodynamic force and gravity, and there is: where f Ax 、f Ay 、f Az are the components of the aerodynamic force on each coordinate axis of this system, and their expressions are respectively: where S ref is the wing reference area; is the dynamic pressure; C X0 、C Xα 、C Xq 、C Yβ 、C Yγ 、C Yp 、C Z0 、C Zα 、C Zq are all aircraft aerodynamic derivatives; where C B / I is the coordinate transformation matrix from the inertial system to the present system, and we have: where C H / I is the transformation matrix from the inertial coordinate system to the track coordinate system; The design of the trajectory law of the position loop of the aircraft using the PID control method includes: Use PID control to control the engine thrust T of the aircraft, and we have: T = K vp (V r - V)+K vI I ev where V r is the reference speed; k vp and K vI are controller gains; is the difference between the reference speed and the aircraft speed; Set the steering method of the aircraft to be achieved by the roll of the aircraft, and use PID control to control the steering of the aircraft. Then the reference signal μ of the aircraft tilt angle μ des is as follows: where μ is the tilt angle of the aircraft; μ des is the reference signal of the tilt angle μ of the aircraft; χ r Reference angle for the aircraft azimuth angle χ given by the azimuth ring; K μp and K μI are controller gains; is the difference between the reference angle of the aircraft azimuth χ and the aircraft azimuth χ; wherein is a rate and amplitude saturation function set for the bank angle μ of the aircraft: where, wherein μ respectively correspond to the maximum and minimum allowable ranges of the aircraft bank angle μ; μ max and μ min are the maximum and minimum values of the allowable aircraft bank angle μ, respectively; μ0 is the μ in the previous control cycle of the controller; d ct is the controller cycle; is the maximum rate limit for μ; Design α using the incremental dynamic inversion method des for design: where α is the angle of attack of the aircraft; α des is the reference signal of the angle of attack α of the aircraft; γ γ Reference angle of the flight path angle γ of the aircraft obtained for the flight path loop; K αp 、K αI 、K αp 、K αI are the controller gains respectively; is the difference between the reference angle of the flight path angle γ of the aircraft and the flight path angle γ of the aircraft; α0 is the angle of attack of the aircraft in the previous control cycle; Δα is the calculated control increment; γ0 is the flight path angle of the aircraft in the previous control cycle; wherein, α0(t) = α(t - d ct ) γ0(t) = γ(t - d ct ) (20) where d ct is the control period; t is the current time step; Among them is a rate and amplitude saturation function set for the angle of attack α of the aircraft: wherein α are respectively the maximum and minimum allowable ranges of the angle of attack α, which are respectively: where α max and α min are the maximum and minimum amplitude limits of the angle of attack α of the aircraft, respectively; is the maximum rate limit for the angle of attack α of the aircraft.

7. A dynamic RCS simulation method for an aircraft during en-route flight according to claim 6, characterized in that: the attitude loop of the aircraft is: wherein is the angular velocity vector of the aircraft body system relative to the inertial system, and its expression is: where p, q, and r are the angular rates of rotation about the three axes of the aircraft's body coordinate system respectively; Among them is the angular velocity vector of the aircraft body system relative to the track coordinate system, and its expression is: wherein is the angular velocity vector of the track coordinate system relative to the inertial coordinate system, and its expression is as follows: where C BH is the transformation matrix for converting the track coordinate system to this system, and its expression is: C H / B is the transformation matrix for coordinate conversion in this system to the track system; design the control law of the attitude loop of the aircraft by using the method of dynamic inversion, including: where p des 、p des 、r des are respectively the reference signals of the angular rates p, q, r rotating about the three axes of the aircraft body system; K μp , K αp , K βp , K μI , K αI , K βI are respectively the controller gains; I eμ ,I eα ,I eβ ,I eλ is the cumulative error in the integrator; μ r is the reference signal for the bank angle μ of the aircraft; α r is the reference signal for the angle of attack α of the aircraft; β r is the reference signal for the sideslip angle β of the aircraft.

8. A method for aircraft dynamic RCS simulation during en-route flight according to claim 7, characterized in that: the angular velocity loop of the aircraft is: L, m, and n are the torques about the x, y, and z axes of the body coordinate system respectively, and their expressions are: l＝l a +l G +l I (31) m = m a +m G +m I (32) n = n a +n G +n I (33) where l a , m a , n a are the aerodynamic moments of the aerodynamic force on each coordinate axis of this system, and their expressions are respectively: Among them, C lβ , C lp , C lr , C m0 , C mα , C mq , C nβ , C np , C nr , are all aircraft aerodynamic derivatives; δ α is the aileron deflection angle of the aircraft; δ r is the rudder deflection angle of the aircraft; δ e is the elevator deflection angle of the aircraft; is the aerodynamic chord length of the aircraft; where l G , m G , n G are the aerodynamic moments of the three axes caused by gravity, and their expressions are respectively: where S x is the x-axis coordinate of the centroid in this system; S y is the y-axis coordinate of the centroid in this system; S z is the z-axis coordinate of the centroid in this system; where l I , m I , n I are the three-axis components of the moment of inertia in this system, and their expressions are respectively: where S = [S x , S y , S z T ; ω = [p, q, r] T ; v is the aircraft velocity vector;​ ρ i is the position vector of the aircraft mass element relative to the origin of the aircraft body system; design the control law of the angular velocity loop of the aircraft by using the method of incremental dynamic inversion, including: where J is the moment of inertia of the aircraft; where p0, q0, and r0 are the values of p, q, and r in the previous control cycle, and there is p0(t) = p(t - d ct ) q0(t) = q(t - d ct ) r0(t) = r(t - d ct ) (39) wherein, where δ a is the aileron deflection angle; δ e is the elevator deflection angle; δ γ is the rudder deflection angle; △δ a is the aileron control increment; △δ e is the elevator control increment; △δ γ is the rudder control increment; δ α0 is the value of the aileron deflection angle in the previous control cycle; δ e0 is the value of the elevator deflection angle in the previous control cycle; δ γ0 is the value of the rudder deflection angle in the previous control cycle, and respectively: where t is the current time step and d ct is the control period.

9. A method for aircraft dynamic RCS simulation during en-route flight according to claim 8, characterized in that: calculate the RCS of the aircraft in different flight postures by using the calculation model of the azimuth angle and altitude angle of the radar relative to the aircraft, and establish a dynamic RCS library, including the following steps: Convert the waypoints of the aircraft and the coordinates of the radar to the coordinates in the earth station geocentric coordinate system, and perform flight path simulation of the aircraft according to the coordinates in the earth station geocentric coordinate system to generate the flight data of the aircraft; Calculate the azimuth angle and altitude angle of the radar in the body coordinate system of the aircraft through coordinate transformation according to the Euler angle and position information of the aircraft in the flight data of the aircraft, wherein the calculation formulas of the azimuth angle and altitude angle of the radar are: where is the azimuth angle of the radar relative to the aircraft, and its value range is [-π, π]; θ' is the altitude angle of the radar relative to the aircraft, and its value range is [0, π]; where (x4, y4, z4) is the coordinate of the radar in the body coordinate system, which is: where (x T , y T , z T ) are the coordinates of the radar in the geocentric coordinate system of the earth station; Q is the transformation matrix from the radar coordinate system to the body coordinate system, which is: where θ is the pitch angle, ψ is the yaw angle, and φ is the roll angle, and they are all the attitude angles of the aircraft; where (x p , y p , z p ) are the coordinates of the trajectory P of the aircraft; Calculate the dynamic RCS value of the aircraft at each moment according to the azimuth angle and altitude angle information of the radar in the body coordinate system of the aircraft to obtain the dynamic RCS library.

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