Aircraft descent arc energy planning method and apparatus

Abstract

This invention provides a method and apparatus for energy planning during the descent phase of an aircraft. After the first-stage engine separates, the minimum RCS attitude of the aircraft is calculated based on its current position and the position of the target detection point. When the aircraft's flight altitude exceeds a preset altitude, its attitude is adjusted to the minimum RCS attitude, and the aircraft is controlled to fly without power. After the aircraft reaches the highest point of its orbit without power, the optimal firing time and optimal firing attitude of the second-stage engine are searched based on the aircraft's tilt angle constraints and reentry point coordinate constraints. The second-stage engine is controlled to ignite at the optimal firing time, and the aircraft's attitude is adjusted to the optimal firing attitude. This achieves the control of the second-stage engine not igniting during the ascent phase of the elliptical orbit. After the aircraft flies without power over the highest point of the elliptical orbit, it flies according to the searched optimal firing time and optimal firing attitude. After the engine is exhausted, it flies according to the attitude constrained by the reentry angle of attack, ensuring that the aircraft's orbit meets the constraints. This increases the difficulty of detection, tracking, and orbit prediction for the aircraft during the second-stage flight phase.

Description

Technical Field

[0001] This application relates to the field of aircraft guidance and control technology, specifically to an energy planning method and device for the arc-drop phase of an aircraft. Background Technology

[0002] The reentry glider combines the orbital characteristics of a reentry vehicle with those of a glider. After the first stage of flight, it can manage energy using a second-stage solid rocket motor, enabling the vehicle to break through the atmosphere in a short time and fly to the reentry point in outer space via a reentry orbit. After entering the reentry point, it pulls up and glides, thus achieving long-distance flight and high-precision landing.

[0003] However, some problems still exist in the second-stage guidance design of reentry gliders. Traditional second-stage guidance designs mainly employ closed-circuit guidance, igniting the engine during the ascent phase of an elliptical orbit and guiding and controlling the engine energy to ensure the reentry point meets requirements such as orbital inclination, position error, and angle of attack. While this design can achieve a certain degree of precise control over the reentry point, it also has significant drawbacks. During the ascent phase of an elliptical orbit, the aircraft generates noticeable infrared and radar signatures, making it easily detectable by surveillance equipment, thus revealing its position and flight trajectory, and making its reentry point relatively predictable.

[0004] Therefore, how to reduce the probability of the aircraft being detected during the engine ignition phase, and improve the aircraft's stealth and trajectory prediction difficulty, while ensuring that the reentry point meets the relevant constraints, is a technical problem that urgently needs to be solved in the field of secondary guidance design for reentry gliders. Summary of the Invention

[0005] This application provides a method and apparatus for energy planning during the descent phase of an aircraft, which can solve the technical problem in the prior art where the aircraft ignites its engine during the ascent phase of an elliptical orbit, and the subsequent flight trajectory is easily detected and tracked.

[0006] In a first aspect, embodiments of this application provide an energy planning method for the arc-drop phase of an aircraft, the energy planning method for the arc-drop phase of an aircraft includes:

[0007] After the first-stage engine of the aircraft separates, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the position of the target detection point.

[0008] When the aircraft's flight altitude exceeds the preset altitude, the aircraft's attitude is adjusted to the lowest RCS attitude, and the aircraft is controlled to fly without power.

[0009] After the spacecraft flies to the highest point of its orbit without power, the optimal firing time and optimal firing attitude of the spacecraft's second-stage engine are searched based on the spacecraft's orbital inclination constraints and reentry point coordinate constraints.

[0010] The system controls the second-stage engine of the aircraft to ignite at the optimal firing moment and adjusts the aircraft's attitude to the optimal firing attitude.

[0011] Secondly, embodiments of this application provide an energy planning device for the landing arc segment of an aircraft, the energy planning device for the landing arc segment of an aircraft comprising:

[0012] The first calculation module is used to calculate the minimum RCS attitude of the aircraft based on the current position of the aircraft and the position of the target detection point after the separation of the first stage engine of the aircraft.

[0013] The first adjustment module is used to adjust the attitude of the aircraft to the lowest RCS attitude and control the aircraft to fly without power when the aircraft's flight altitude is greater than the preset altitude.

[0014] The second calculation module is used to search for the optimal firing time and optimal firing attitude of the second-stage engine of the aircraft after the aircraft flies to the highest point of the orbit without power, based on the orbital inclination constraints and reentry point coordinate constraints of the aircraft.

[0015] The second adjustment module is used to control the second-stage engine of the aircraft to ignite at the optimal firing moment and adjust the attitude of the aircraft to the optimal firing attitude.

[0016] The beneficial effects of the technical solutions provided in this application include:

[0017] After the first-stage engine separates from the spacecraft, the minimum RCS attitude of the spacecraft is calculated based on its current position and the position of the target detection point. When the spacecraft's flight altitude exceeds a preset altitude, its attitude is adjusted to the minimum RCS attitude, and the spacecraft is controlled to fly without power. After the spacecraft flies without power to the highest point of its orbit, the optimal firing time and optimal firing attitude of the spacecraft's second-stage engine are searched based on the orbital inclination and coordinate constraints of the reentry point. The spacecraft's second-stage engine is controlled to ignite at the optimal firing time, and the spacecraft's attitude is adjusted to the optimal firing attitude. This achieves the following: after the first-stage engine burns out and separates, the second-stage engine does not ignite during the ascent phase of the elliptical orbit, but instead rises to a certain altitude and adjusts to the minimum RCS attitude for the target detection point. After flying without power over the highest point of the elliptical orbit, it flies according to the searched optimal firing time and optimal firing attitude. After the engine exhausts, it flies according to the attitude constrained by the reentry angle of attack, ensuring that the spacecraft's orbit meets the constraints, thereby increasing the difficulty of detection, tracking, and orbit prediction of the spacecraft during the second-stage flight phase. Attached Figure Description

[0018] Figure 1 This is a flowchart illustrating an embodiment of the energy planning method for the landing arc segment of an aircraft according to this application;

[0019] Figure 2 A schematic diagram illustrating the division of the reentry glide trajectory phase for an aircraft.

[0020] Figure 3 This is a schematic diagram showing the relationship between the combustion process of the second-stage engine and the pitch angle of the aircraft.

[0021] Figure 4 This is a functional module diagram of an embodiment of the energy planning device for the landing arc segment of the aircraft in this application. Detailed Implementation

[0022] To enable those skilled in the art to better understand the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present application, and not all embodiments. Based on the embodiments of the present application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present application.

[0023] First, some of the technical terms used in this application will be explained to help those skilled in the art understand this application.

[0024] Lowest RCS attitude: refers to the attitude of an aircraft during flight that minimizes its radar cross section (RCS) by adjusting its own attitude.

[0025] To make the objectives, technical solutions, and advantages of this application clearer, the embodiments of this application will be described in further detail below with reference to the accompanying drawings.

[0026] In a first aspect, embodiments of this application provide an energy planning method for the arc-drop segment of an aircraft.

[0027] In one embodiment, please refer to the following: Figure 1 and Figure 2 , Figure 1 This is a flowchart illustrating the first embodiment of the energy planning method for the arc-drop segment of the aircraft in this application. Figure 1 As shown, the energy planning method for the descent segment of an aircraft includes:

[0028] Step S1: After the first-stage engine of the aircraft separates, calculate the minimum RCS attitude of the aircraft based on the current position of the aircraft and the position of the target detection point.

[0029] In one embodiment, step S1 specifically includes the following steps:

[0030] Step S101: During flight, the guidance system determines the current vector position of the aircraft in the launch inertial frame. Vector position of the target detection point in the launch inertial frame Determine the vector of the aircraft relative to the target detection point:

[0031]

[0032] in, This is the vector of the aircraft relative to the detection point.

[0033] Step S102: Normalize the vector of the aircraft relative to the target detection point to obtain the corresponding first unit vector.

[0034] The magnitude of the vector of the aircraft relative to the target detection point is:

[0035]

[0036] in, for The model, for The component in the X direction, for The component in the Y direction, for The component in the Z direction.

[0037] right Perform vector normalization to obtain the first unit vector;

[0038]

[0039] in, It is the first unit vector.

[0040] Step S103: Normalize the current vector position of the aircraft to obtain the corresponding second unit vector.

[0041] The modulus of the aircraft's current vector position is:

[0042]

[0043] in, for The model, for The component in the X direction, for The component in the Y direction, for The component in the Z direction.

[0044] right By performing vector normalization, we obtain the corresponding second unit vector:

[0045]

[0046] in, It is the second unit vector.

[0047] Step S104: Calculate the cross product of the first unit vector and the second unit vector in the opposite direction to obtain the reference axis:

[0048]

[0049] in, Used as the reference axis.

[0050] Step S105: Perform a cross product calculation on the reference axis and the first unit vector to obtain the corrected reference axis:

[0051]

[0052] in, This is the corrected reference axis.

[0053] Step S106: Based on the first unit vector, the corrected reference axis, and the reference axis, calculate the minimum RCS pitch angle, minimum RCS yaw angle, and minimum RCS roll angle in the minimum RCS attitude:

[0054]

[0055] ψ = a sin(-Zx)

[0056] γ = a tan 2(Zy, Zz)

[0057] in, ψ is the lowest RCS pitch angle, γ is the lowest RCS yaw angle, γ is the lowest RCS roll angle, Xx is the x-component of the first unit vector, Yx is the x-component of the corrected reference axis, Zx is the x-component of the reference axis, Zy is the y-component of the reference axis, and Zz is the z-component of the reference axis.

[0058] In an alternative implementation, matrix A can be constructed based on the first unit vector, the corrected reference axis, and the reference axis:

[0059]

[0060] The minimum RCS pitch angle, minimum RCS yaw angle, and minimum RCS roll angle are calculated based on matrix A.

[0061]

[0062] ψ = a sin(-A13)

[0063] γ = a tan 2(A23,A33)

[0064] Where A11 is the element Xx in the first row and first column of matrix A, A12 is the element Yx in the first row and second column of matrix A, A13 is the element Zx in the first row and third column of matrix A, A23 is the element Zy in the second row and third column of matrix A, and A33 is the element Zz in the third row and third column of matrix A.

[0065] As a preferred embodiment, after the first-stage engine of the aircraft separates, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the position of the target detection point, including: when a new target detection point position is received, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the position of the new target detection point.

[0066] Explanatory purposes, the detection point in this embodiment can be a radar detection point. During flight, the aircraft may encounter multiple detection points. One of these points can be selected as the target detection point. Furthermore, the system supports updating the target detection point and its position during flight to calculate the aircraft's minimum RCS attitude relative to the target detection point in real time, thereby increasing the difficulty of detecting and tracking the aircraft.

[0067] Step S2: When the aircraft's flight altitude is greater than the preset altitude, adjust the aircraft's attitude to the lowest RCS attitude and control the aircraft to fly without power.

[0068] It is worth noting that the preset altitude in this embodiment can be set to 70KM or 100KM. By setting the second-stage engine to not ignite during the elliptical orbit ascent segment after the primary active stage engine has completely burned out and separated, and instead ascending to the preset altitude and adjusting to the lowest RCS attitude for the detection point, this invention can improve the stealth of the aircraft and increase the difficulty of its detection and tracking.

[0069] Step S3: After the aircraft flies to the highest point of the orbit without power, search for the optimal firing time and optimal firing attitude of the second-stage engine of the aircraft based on the orbital inclination angle constraint and the reentry point coordinate constraint of the aircraft.

[0070] In one embodiment, before searching for the optimal firing time and optimal firing attitude of the second-stage engine based on the orbital inclination constraint and reentry point coordinate constraint of the aircraft's reentry point, the method further includes:

[0071] like Figure 3As shown, after the second-stage engine ignites, the correspondence between the second-stage engine combustion process and the aircraft's pitch angle is as follows: the time from second-stage engine ignition to the completion of combustion is the engine combustion time (Tfire). This combustion time (Tfire) is divided into three segments: the first third of the second-stage engine combustion period, from 0 to 0.333 * Tfire, is the attitude stabilization phase, where the aircraft's pitch angle remains at the initial pitch angle at the time of engine ignition; the middle third of the second-stage engine combustion period, from 0.333 * Tfire to 0.666 * Tfire, is the alternating attitude phase, where the aircraft's pitch angle linearly adjusts from the initial pitch angle to its final value; and the last third of the second-stage engine combustion period, from 0.666 * Tfire to Tfire, is the attitude maintenance phase, where the aircraft's pitch angle remains at its final value. Therefore, as... Figure 3 As shown, based on the correspondence between the second-stage engine combustion process and the aircraft pitch angle, the corresponding aircraft pitch angle can be obtained by time interpolation.

[0072] Specifically, step S3 searches for the optimal firing time and optimal firing attitude of the second-stage engine based on the orbital inclination constraint and the reentry point coordinate constraint of the aircraft. This includes the following steps: searching for the optimal firing time of the second-stage engine based on the orbital inclination constraint of the aircraft; searching for the optimal yaw angle of the aircraft after the second-stage engine ignition based on the Z-axis coordinate constraint of the aircraft's reentry point; and searching for the optimal final pitch angle of the aircraft after the second-stage engine ignition based on the X-axis coordinate constraint of the aircraft's reentry point.

[0073] In one embodiment, searching for the optimal firing time of the second-stage engine of the aircraft based on the aircraft's orbital inclination constraints includes:

[0074] Step S311: Starting from the current time T, the aircraft performs unpowered external thrust to the planned ignition time Tign for virtual ignition. Based on the aircraft's position and velocity when the second-stage engine performs virtual ignition at the planned ignition time, and the aircraft's second-stage flight pitch angle, second-stage flight yaw angle, and second-stage flight roll angle after virtual ignition, the aircraft is calculated to perform powered external thrust until the second-stage engine burns out, then switches to unpowered external thrust, until the aircraft is pushed to the reentry point altitude Hend, corresponding to the first inclination angle θzr1 of the orbit.

[0075] The second-stage flight pitch angle is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle. In this correspondence, the final pitch angle value is the planned final pitch angle value, which is the final pitch angle value designed in the aircraft's standard orbit. The second-stage flight yaw angle maintains the yaw angle ψ at the virtual ignition moment, and the second-stage flight roll angle is zero. In the first iteration, the planned ignition moment Tign is the designed second-stage engine ignition moment in the aircraft's standard orbit.

[0076] Step S312: Calculate the inclination error between the first inclination angle of the track and the preset standard inclination angle value:

[0077] dθzr=θzr1-θbz

[0078] Where dθzr is the tilt error, θzr1 is the first tilt angle of the orbit, and θbz is the standard tilt angle value. The standard tilt angle value is the tilt angle of the reentry point orbit designed in the standard orbit of the spacecraft.

[0079] Step S313: Add the preset time adjustment step dT to the planned ignition time Tign to obtain the adjusted ignition time Tign+dT. Starting from the current time T, the aircraft performs unpowered external push to the adjusted ignition time Tign+dT for virtual ignition. Based on the aircraft's position and speed when the second-stage engine performs virtual ignition at the adjusted ignition time, and the aircraft's second-stage flight pitch angle, second-stage flight yaw angle, and second-stage flight roll angle after virtual ignition, calculate the aircraft's powered external push until the second-stage engine burns out, then switch to unpowered external push, until the aircraft is pushed to the reentry point altitude Hend, corresponding to the second orbital inclination angle θzr2.

[0080] The methods for obtaining the secondary flight pitch angle, secondary flight yaw angle, and secondary flight roll angle are the same as those for obtaining them in step S311, and will not be repeated here.

[0081] Step S314: Calculate the derivative of the orbit inclination angle relative to the time adjustment step size based on the first and second orbit inclination angles.

[0082] dθzrdt=(θzr2-θzr11) / dT

[0083] Where θzr1 is the first inclination angle of the orbit, θzr2 is the second inclination angle of the orbit, dθzrdt is the derivative of the orbit inclination angle with respect to the time adjustment step, and dT is the time adjustment step.

[0084] Step S315: Based on the time adjustment step size and the ratio of the inclination error to the derivative of the orbital inclination angle with respect to the time adjustment step size, update the planned ignition time.

[0085] Tign=Tign-dT-dθzr / dθzrdt

[0086] After updating the planned ignition time, the cycle of step S311 and step S315 continues until the tilt error is less than the preset tilt error threshold or the preset number of iterations is reached. The virtual ignition time corresponding to the end of the iteration is taken as the optimal ignition time.

[0087] As an example, the tilt angle error threshold in this embodiment is 0.1°, and the preset number of iterations is 10, that is, until the number of iterations is greater than 10 or dθzr is calculated to be less than 0.1°. The virtual ignition time obtained at this point is the optimal ignition time Tign that satisfies the reentry point orbit tilt angle constraint. opt .

[0088] In one embodiment, based on the Z-axis coordinate constraint of the aircraft's reentry point, the search for the optimal yaw angle of the aircraft after the ignition of the second-stage engine includes:

[0089] Step S321: Starting from the current time T, the spacecraft performs unpowered extrapolation to the optimal launch time Tign. opt Virtual ignition is performed. Based on the position and speed of the aircraft when the second-stage engine is ignited at its optimal ignition moment, the second-stage flight pitch angle, second-stage flight yaw angle, and second-stage flight roll angle of the aircraft after virtual ignition are calculated. The aircraft is then pushed outward with power until the second-stage engine is exhausted and then switched to unpowered pushout, until it is pushed outward to the reentry point altitude Hend, corresponding to the first Z-axis coordinate Z1.

[0090] The second-stage pitch angle of the aircraft after virtual ignition is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle. At this time, the final pitch angle in the correspondence between the second-stage engine combustion process and the pitch angle is the planned final pitch angle, which is the final pitch angle designed in the standard trajectory of the aircraft. The second-stage yaw angle during the first iteration maintains the yaw angle ψ at the optimal ignition time, and the second-stage roll angle is zero.

[0091] Step S322: Calculate the first Z-axis error between the first Z-axis coordinate and the preset standard Z-axis coordinate.

[0092] dZ1=Z1-Zbz

[0093] Where dZ1 is the first Z-axis error, Z1 is the first Z-axis coordinate, and Zbz is the standard Z-axis coordinate. The standard Z-axis coordinate is the Z-axis coordinate of the reentry point designed in the standard orbit of the spacecraft.

[0094] Step S323: Add the preset yaw adjustment step dψ to the second-stage flight yaw angle ψ to obtain the adjusted yaw angle ψ+dψ. Based on the position and speed of the aircraft when the second-stage engine is virtually ignited at the optimal moment, the second-stage flight pitch angle, adjusted yaw angle ψ+dψ, and second-stage flight roll angle of the aircraft after virtual ignition, calculate the second Z-coordinate Z2 corresponding to the time when the aircraft is pushed outward with power until the second-stage engine is exhausted and then switches to unpowered pushout, until the aircraft is pushed outward to the reentry point altitude Hend.

[0095] The methods for obtaining the secondary flight pitch angle and secondary flight roll angle are the same as those for obtaining them in step S321, and will not be repeated here.

[0096] Step S324: Calculate the second Z-axis error between the second Z-axis coordinate and the standard Z-axis coordinate.

[0097] dZ2=Z2-Zbz

[0098] Where dZ2 is the second Z-axis error and Z2 is the second Z-axis coordinate.

[0099] Step S325: Based on the first Z-axis error and the second Z-axis error, calculate the derivative of the Z-axis error with respect to the yaw angle adjustment step size:

[0100] dZdψ=(dZ2-dZ1) / dψ

[0101] Where dZdψ is the derivative of the Z-axis error with respect to the yaw angle adjustment step size, dZ1 is the first Z-axis error, and dZ2 is the second Z-axis error.

[0102] Step S326: Update the planned yaw angle based on the yaw angle adjustment step size and the ratio of the Z-axis error to the derivative of the Z-axis error relative to the yaw angle adjustment step size.

[0103] ψ=ψ-dψ-dZ / dZdψ

[0104] After updating the planned yaw angle, continue to iterate through step S321 and step S326 until the Z-axis error is less than the preset Z-axis error threshold or the preset number of iterations is reached. The yaw angle after virtual ignition at the end of the iteration is taken as the optimal yaw angle.

[0105] As an example, in this embodiment, the Z-axis error threshold is 500 meters, and the preset number of iterations is 10, that is, until the number of iterations is greater than 10 or the calculated dZ is less than 500 meters. At this point, the optimal yaw angle ψ that satisfies the reentry point Z-axis error constraint after secondary ignition is obtained. opt .

[0106] In one embodiment, based on the X-axis coordinate constraint of the aircraft's reentry point, the search for the optimal final pitch angle of the aircraft after the second-stage engine ignition includes:

[0107] Step S331: Starting from the current time T, the spacecraft performs unpowered extrapolation to the optimal launch time Tign. opt A virtual ignition is performed, based on the aircraft's position and velocity at the optimal ignition moment of the second-stage engine. The corresponding second-stage flight pitch angle ψ of the aircraft after the virtual ignition is determined. opt Optimal yaw angle and roll angle ψ opt The calculation is performed to determine the first X-axis coordinate X1 corresponding to the time when the aircraft pushes outward with power until the second-stage engine burns out and then switches to unpowered pushout, until it pushes outward to the reentry point altitude Hend.

[0108] The second-stage pitch angle of the aircraft after virtual ignition is obtained by interpolation over time based on the correspondence between the second-stage engine combustion process and the pitch angle. In the first iteration, the final pitch angle value in the correspondence between the second-stage engine combustion process and the pitch angle is the planned final pitch angle value, which is the final pitch angle value designed in the standard trajectory of the aircraft. The second-stage roll angle is zero.

[0109] Step S332: Calculate the first X-axis error between the first X-axis coordinate and the preset standard X-axis coordinate.

[0110] dX1=X1-Xbz

[0111] Where dX1 is the first X-axis error, X1 is the first X-axis coordinate, and Xbz is the standard X-axis coordinate. The standard X-axis coordinate is the designed X-axis coordinate of the reentry point in the standard orbit of the spacecraft.

[0112] Step S333: Set the final value of the planned pitch angle. Add the preset pitch angle final value adjustment step size Obtain the final value of the adjusted pitch angle Based on the aircraft's position and velocity at the optimal ignition moment of the second-stage engine during virtual ignition, the corresponding second-stage flight pitch angle and optimal yaw angle ψ of the aircraft after virtual ignition are determined. opt And the second-stage roll angle, calculate the second X-axis coordinate X2 corresponding to the time when the aircraft pushes outward with power until the second-stage engine is exhausted and then switches to unpowered pushout, until it pushes outward to the reentry point altitude Hend.

[0113] Step S334: Calculate the second X-axis error between the second X-axis coordinate and the standard X-axis coordinate.

[0114] dX2=X2-Xbz

[0115] Where dX2 is the second X error.

[0116] Step S335: Based on the first X-axis error and the second X-axis error, calculate the derivative of the adjustment step size of the X-axis error relative to the final value of the pitch angle:

[0117]

[0118] in, The derivative of the pitch error relative to the final value of the pitch angle adjustment step size, dX1 is the first X-axis error, and dX2 is the second X-axis error.

[0119] Step S336: Adjust the step size based on the final pitch angle value, and update the planned final pitch angle value by using the ratio of the X-axis error to the derivative of the X-axis error relative to the final pitch angle value adjustment step size.

[0120]

[0121] After updating the final value of the planned pitch angle, continue to loop step S331 and step S336 and iterate until the X-axis error is less than the preset X-axis error threshold or the preset number of iterations is reached. The final value of the yaw angle after virtual ignition at the end of the iteration is taken as the optimal final value of the pitch angle.

[0122] As an example, in this embodiment, the X-axis error threshold is 500 meters, and the preset number of iterations is 10, i.e., until the number of iterations is greater than 10 or the calculated dX is less than 500 meters. The final optimal pitch angle obtained at this point satisfies the X-axis error constraint at the reentry point after secondary ignition.

[0123] In one embodiment, the algorithm for determining the position and velocity of the aircraft during its unpowered outward glide to the virtual ignition moment is as follows.

[0124] Step S341: Calculate the gravitational acceleration Gxyz of the spacecraft at the current moment. k Initial time k = 0, t k =0;

[0125] Step S342: Set the extrapolation step size to 0.02s;

[0126] Step S343: Calculate the position result X for the next extrapolation step. k+1 ,Y k+1 Z k+1 :

[0127]

[0128] Step S344: Calculate the acceleration Gxyz at the current position based on the updated position result. k+1 Calculate the speed result of the next extrapolation:

[0129]

[0130] Step S345: Update the current iteration time t k+1 = t k + step, when t k > Tign, the unpowered extrapolation ends, and the extrapolation result is the position and velocity at the moment of gliding to Tign without power.

[0131] Step S346: Calculate the absolute height corresponding to the current position. When the absolute height Habs < Hend, the unpowered extrapolation ends, and the extrapolation result is the position and velocity at the moment of gliding to the target height Tend without power.

[0132] In one embodiment, after the secondary engine of the aircraft ignites, it is extrapolated with power until the moment when the secondary engine combustion is completed. The algorithms for the flight time, speed, and position of the aircraft are as follows:

[0133] Step S351: Calculate the gravitational acceleration Gxyz of the aircraft at the current moment k , at the initial moment k = 0, t k = 0;

[0134] Step S352: Set the extrapolation step size step = 0.02 s;

[0135] Step S353: Assume ignition at the initial moment. According to the apparent acceleration provided by the engine and the total combustion duration t endfire , t k Interpolate with endpoint limiting to obtain the apparent acceleration a at the current moment k and the apparent acceleration a at the next moment k+1 , then:

[0136] a = 0.5·(a k + a k+1 )

[0137] Step S354: Project the apparent acceleration a onto the current initial attitude and calculate the resulting apparent acceleration dVxyz:

[0138]

[0139] Step S355: Calculate the position results X k+1 , Y k+1 , Z k+1 :

[0140]

[0141] Step S356: Calculate the acceleration Gxyz at the current position according to the updated position result, and calculate the velocity result of the next extrapolation: k+1 , calculate the velocity result of the next extrapolation:

[0142]

[0143] Step S357: Update the current iteration time t k+1 =t k +step, when t k+1 >t endfire At that time, the extrapolation ends, and the flight time T at the moment when the second-stage engine exhausts is obtained. endwt =t k +T0+T start Position and velocity XYZ endwt Vxyz endwt .

[0144] Step S4: Control the second-stage engine of the aircraft to ignite at the optimal firing moment, and adjust the attitude of the aircraft to the optimal firing attitude.

[0145] In a preferred embodiment, the method further includes:

[0146] After the second-stage engine of the aircraft has completely burned out and the aircraft's altitude is lower than the preset altitude, the attitude is adjusted to the reentry angle of attack α0 based on the current position and velocity. The calculation steps are as follows:

[0147] Determine the vector of the spacecraft relative to the Earth's center:

[0148]

[0149] in, Let be the vector of the spacecraft relative to the Earth's center, and Rox, Roy, and Roz be the three components of the Earth's center in the launch inertial frame, respectively.

[0150] Calculate the vector of the aircraft relative to the Earth's center and the velocity vector of the aircraft in the launch inertial frame. dot product:

[0151] Dxyzp=XpVx+YpVy+ZpVz

[0152] Where Dxyzp is the dot product, and Vx, Vy, and Vz are the three components of the velocity vector of the spacecraft in the launch inertial frame, respectively.

[0153] Determine the magnitude of the spacecraft's vector relative to the Earth's center and the magnitude of the spacecraft's velocity vector:

[0154]

[0155] Where Nxyz is the magnitude of the vector of the aircraft relative to the Earth's center, and NVxyz is the magnitude of the velocity vector of the aircraft.

[0156] Based on the dot product, the magnitude of the aircraft's vector relative to the Earth's center, and the magnitude of the aircraft's velocity vector, calculate the reentry pitch angle, reentry yaw angle, and reentry roll angle when the reentry angle of attack satisfies the constrained angle of attack:

[0157]

[0158] γzr=0

[0159] in, ψzr is the reentry pitch angle, γzr is the reentry yaw angle, γzr is the reentry roll angle, Dxyzp is the dot product, Nxyz is the magnitude of the vector of the spacecraft relative to the Earth's center, and NVxyz is the magnitude of the spacecraft's velocity vector.

[0160] It is worth noting that this application, by controlling the aircraft to fly over the highest point of the unpowered trajectory, employs virtual ignition and open-loop prediction methods to plan appropriate ignition timing and energy dissipation attitude, ensuring constraints on trajectory inclination and reentry point errors. This method effectively increases the difficulty of detection, tracking, and trajectory prediction for the reentry-glide vehicle in the second-stage flight phase, while meeting the required constraints for entering the gliding phase.

[0161] In a specific embodiment, assume that the current position of an aircraft is (629312; 148365; 11404m); the position of the target radar observation station is (1064700; -60478; -90723m); the current pitch angle of the aircraft is -28.8°, the yaw angle is 30.8°, and the roll angle is -0.55°.

[0162] After reaching an altitude of 70 km, the minimum RCS attitude is calculated based on the aircraft's current position and the radar station's position. First, the reference axis Xx is calculated as (0.88; -0.42 -0.21); Yref is calculated as (0.97; 0.23 0.02); Zz is calculated as (-0.04; 0.22; -0.61); and the corrected reference Y-axis Yy is calculated as (-0.30; -0.55; -0.17). The Xx, Yy, and Zz are combined to form matrix A, and the inverse solution yields the minimum RCS attitude as follows: pitch angle -19.06°; yaw angle 2.29°; and roll angle 160.54°. In this attitude, the aircraft's nose is directly facing the radar station, minimizing the RCS value detected by the radar station.

[0163] In a specific embodiment, assume that the current time of an aircraft is 360s; its current position is (629312; 148365; 11404m); its current velocity is (2005; -954; 65m / s) and it has already passed the highest point of its orbit; the aircraft's current pitch angle is -28.8°; yaw angle is 30.8°; and roll angle is -0.55°. The standard orbit design has a planned second-stage engine ignition time of 388.012s, and the planned final pitch angle after the second-stage engine burns out is 89°. The axial apparent acceleration curve of the aircraft after the second-stage engine ignition has been obtained in advance.

[0164] The standard orbital design reentry point coordinates are (914951; -3062; -39372), the required reentry point orbital inclination angle is -15.8° ± 1°, and the reentry point is maintained at a 0 angle of attack. Energy planning for the reduced arc segment is then performed accordingly.

[0165] Firstly, as Figure 3 As shown, the combustion time of the second-stage engine is divided into three equal parts: the attitude stabilization segment before ignition, the alternating attitude segment during the middle of ignition, and the attitude maintenance segment after ignition. After ignition, the corresponding pitch angle attitude is obtained by interpolation of the engine combustion time.

[0166] Then, the ignition time is optimized to obtain the optimal ignition time, and a loop is constructed. Based on the current position and speed, the virtual ignition is performed by extrapolating without power to 388.012s. Then, the virtual ignition is performed by extrapolating with power. During this period, the pitch angle of the second-stage flight is obtained by interpolation based on the combustion time of the second-stage engine after ignition, the yaw angle of the second-stage flight remains unchanged, and the roll angle of the second-stage flight is 0. Extrapolation with power is performed. When the second-stage engine combustion is exhausted, the virtual extrapolation continues until the Y-axis is less than or equal to the Y-axis of the reentry point (-3062m) and then stops. The orbital inclination angle of the reentry point is calculated to be -16.104°.

[0167] Then, the time adjustment step was set to 0.05s. Based on the current position and speed, the unpowered extrapolation was performed until ignition at 388.062s. Then, the powered extrapolation was performed. During this period, the pitch angle of the second-stage flight was obtained by interpolation based on the combustion time of the second-stage engine after ignition. The yaw angle of the second-stage flight remained unchanged, and the roll angle of the second-stage flight was 0. Powered extrapolation was then performed. When the second-stage engine combustion was exhausted, the unpowered extrapolation was continued until the Y-axis was less than or equal to the Y-axis of the reentry point (-3062m) and then stopped. The orbital inclination angle of the reentry point was calculated to be -16.098°.

[0168] The calculated value is dθzrdt = 0.1122; the ignition time for the next cycle is 390.653s.

[0169] When the number of iterations is greater than 10, or the reentry inclination angle deviation from the standard orbit constraint is less than 0.1°, the loop is exited, and the optimal firing time of 390.653s is obtained; the number of iterations is 1.

[0170] Furthermore, the yaw angle is optimized to obtain the optimal yaw angle, and a loop is constructed. Based on the current position and speed, the unpowered extrapolation is performed until the optimal ignition time of 390.653s. Then, powered extrapolation is performed, during which the second-stage flight pitch angle is obtained by interpolation based on the second-stage engine combustion time after ignition. The second-stage flight yaw angle is 30.8°, and the second-stage flight roll angle is 0. Powered extrapolation continues until the second-stage engine combustion is exhausted, at which point unpowered extrapolation stops when the Y-axis is less than or equal to the reentry point Y-axis (-3062m). The calculated Z-axis position deviation of the reentry point is 4296m.

[0171] Then, the yaw angle adjustment step was designed to be 0.1°. Based on the current position and speed, the engine was pushed outward without power until ignition at 390.653s. Then, it was pushed outward with power. During this period, the second-stage flight pitch angle was obtained by interpolation based on the second-stage engine combustion time after ignition. The yaw angle was adjusted to 30.9°, and the second-stage flight roll angle was 0. Powered pushout was then performed. After the second-stage engine combustion was exhausted, the engine was pushed outward without power until the Y-axis was less than or equal to the Y-axis of the reentry point (-3062m) and then stopped. The Z-axis position deviation of the reentry point was calculated to be 4153m.

[0172] The calculation yields dZdψ = -81837.5; the second-order yaw angle for the next cycle is 32.907°.

[0173] When the number of iterations is greater than 10, or the deviation of the Z-axis reentry point from the standard orbit constraint is less than 500m, the loop is exited, and the optimal yaw angle is obtained as 32.907°; the number of iterations is 1.

[0174] Furthermore, the final pitch angle value is optimized to obtain the optimal final pitch angle value, and a loop is constructed. Based on the current position and speed, the unpowered extrapolation proceeds to the optimal ignition time of 390.653s. Then, powered extrapolation continues, during which the second-stage flight pitch angle is obtained by interpolation based on the second-stage engine combustion time after ignition, with an optimal yaw angle of 32.907° and a second-stage flight roll angle of 0, followed by powered extrapolation. When the second-stage engine combustion is exhausted, unpowered extrapolation continues until the Y-axis is less than or equal to the reentry point Y-axis (-3062m), at which point the loop stops. The calculated X-axis position deviation of the reentry point is -1297m.

[0175] Then, the final pitch angle adjustment step was set to 0.1°, and the total pitch angle after the second-stage engine combustion was exhausted was set to 89.1°. The pitch angle was then regenerated as follows. Figure 3The table shows the pitch angle attitude list interpolated with the engine combustion time; then, based on the current position and speed, the engine is pushed outward without power until ignition at 390.653s; then, it is pushed outward with power, during which the pitch angle is obtained by interpolation based on the combustion time of the second-stage engine after ignition. After the combustion time is entered, the optimal yaw angle is 32.907° and the second-stage roll angle is 0, and then the engine is pushed outward with power; when the second-stage engine combustion is exhausted, the engine is pushed outward without power until the Y-axis is less than or equal to the Y-axis of the reentry point (-3062m) and then it stops. The X-axis position deviation of the reentry point is calculated to be -1257m.

[0176] The calculations have been completed to date. After the second-stage engine burns out in the next cycle, the pitch angle will be 82.76°.

[0177] When the number of iterations is greater than 10, or the deviation of the X-axis reentry point from the standard orbit constraint is less than 600m, the loop is exited, and the optimal pitch angle attitude after the second-stage engine burns out is obtained as 82.76°; the number of iterations is 1.

[0178] When the flight time reaches the optimal ignition time of 390.653s, the second-stage ignition is performed. Subsequently, the flight follows the generated optimal pitch final value, optimal yaw angle, and roll angle until the reentry point. The actual reentry error is (-543; 2.83; 1862m); the reentry point tilt angle error is 0.08°.

[0179] In a specific embodiment, assume that the current time of an aircraft is 360s; the current position is (927385; -7599; -42231m); the current speed is (2796; -1242; -1001m / s); the second-stage engine has been exhausted; and it needs to re-enter at an angle of attack of α0 = 0 at the re-entry point. Calculate the 0 angle-of-attack re-entry attitude accordingly.

[0180] The calculation yielded:

[0181] Dxyzp=2645491331.02;

[0182] Nxyz = 928377.99

[0183] NVxyz = 3220.15

[0184] The reentry attitude at 0 angle of attack is as follows: reentry pitch angle is 62.24°, reentry yaw angle is 18.11°, and reentry roll angle is 0°.

[0185] Secondly, embodiments of this application also provide an energy planning device for the landing arc segment of an aircraft.

[0186] In one embodiment, reference is made to Figure 4 , Figure 4This is a functional module diagram of an embodiment of the energy planning device for the landing arc segment of an aircraft according to this application. Figure 4 As shown, the energy planning device for the landing arc segment of the aircraft includes:

[0187] The first calculation module is used to calculate the minimum RCS attitude of the aircraft based on the current position of the aircraft and the position of the target detection point after the separation of the first stage engine of the aircraft.

[0188] The first adjustment module is used to adjust the attitude of the aircraft to the lowest RCS attitude and control the aircraft to fly without power when the aircraft's flight altitude is greater than the preset altitude.

[0189] The second calculation module is used to search for the optimal firing time and optimal firing attitude of the second-stage engine of the aircraft after the aircraft flies to the highest point of the orbit without power, based on the orbital inclination angle constraint and the reentry point coordinate constraint of the aircraft.

[0190] The second adjustment module is used to control the second-stage engine of the aircraft to ignite at the optimal firing moment and adjust the attitude of the aircraft to the optimal firing attitude.

[0191] Furthermore, in one embodiment, the first computing module is further configured to:

[0192] Based on the current vector position of the aircraft in the launch inertial frame and the vector position of the target detection point in the launch inertial frame, determine the vector of the aircraft relative to the target detection point;

[0193] The vector of the aircraft relative to the target detection point is vector normalized to obtain the corresponding first unit vector;

[0194] The current vector position of the aircraft is vector normalized to obtain the corresponding second unit vector;

[0195] The reference axis is obtained by performing the cross product of the first unit vector and the second unit vector in the opposite direction.

[0196] The cross product of the reference axis and the first unit vector is calculated to obtain the corrected reference axis.

[0197] Based on the first unit vector, the corrected reference axis, and the reference axis, calculate the minimum RCS pitch angle, minimum RCS yaw angle, and minimum RCS roll angle in the minimum RCS attitude:

[0198]

[0199] ψ = a sin(-Zx)

[0200] γ = a tan 2(Zy, Zz)

[0201] in, ψ is the lowest RCS pitch angle, γ is the lowest RCS yaw angle, γ is the lowest RCS roll angle, Xx is the x-component of the first unit vector, Yx is the x-component of the corrected reference axis, Zx is the x-component of the reference axis, Zy is the y-component of the reference axis, and Zz is the z-component of the reference axis.

[0202] Furthermore, in one embodiment, the second computing module is also used for:

[0203] Based on the orbital inclination constraints of the spacecraft, search for the optimal firing time of the spacecraft's second-stage engine;

[0204] Based on the Z-axis coordinate constraint of the aircraft's reentry point, search for the optimal yaw angle of the aircraft after the second-stage engine ignition;

[0205] Based on the X-axis coordinate constraint of the aircraft's reentry point, search for the final value of the aircraft's optimal pitch angle after the second-stage engine ignites.

[0206] Furthermore, in one embodiment, the second computing module is also used for:

[0207] After the second-stage engine of the aircraft is set to ignite, the correspondence between the combustion process of the second-stage engine and the pitch angle of the aircraft is as follows:

[0208] For the first third of the second-stage engine's combustion period, the aircraft's pitch angle remains the initial pitch angle at the time of the second-stage engine's ignition.

[0209] During the middle third of the second-stage engine combustion period, the aircraft pitch angle is linearly adjusted from the initial pitch angle to the final pitch angle value.

[0210] During the last third of the second-stage engine's combustion period, the aircraft maintains its final pitch angle value.

[0211] Furthermore, in one embodiment, the second calculation module is also used to: cyclically calculate the first inclination angle of the trajectory when the aircraft is pushed out to the reentry point altitude based on the aircraft's position and velocity when the second-stage engine performs virtual ignition at the planned ignition time, the second-stage flight pitch angle, the second-stage flight yaw angle, and the second-stage flight roll angle after virtual ignition, wherein the second-stage flight pitch angle is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle, the second-stage flight yaw angle is the yaw angle at the virtual ignition time, and the second-stage flight roll angle is zero;

[0212] Calculate the inclination error between the first inclination angle of the track and the preset standard inclination angle value;

[0213] The adjusted ignition time is obtained by adding the preset time adjustment step to the planned ignition time. Based on the position and speed of the aircraft when the second-stage engine is virtually ignited at the adjusted ignition time, and the second-stage flight pitch angle, second-stage flight yaw angle and second-stage flight roll angle of the aircraft after virtual ignition, the second inclination angle of the trajectory when the aircraft is pushed out to the reentry point altitude is calculated.

[0214] Calculate the derivative of the orbit inclination angle relative to the time adjustment step based on the first and second orbit inclination angles;

[0215] The planned ignition time is updated based on the time adjustment step size and the ratio of the inclination error to the derivative of the orbital inclination with respect to the time adjustment step size. This process is iterated until the inclination error is less than the preset inclination error threshold or the preset number of iterations is reached. The virtual ignition time corresponding to the end of the iteration is taken as the optimal ignition time.

[0216] Furthermore, in one embodiment, the second calculation module is also used to: cyclically calculate the first Z-axis coordinates of the aircraft when it is pushed to the reentry point altitude based on the aircraft's position and velocity when the second-stage engine performs virtual ignition at the optimal ignition time, the second-stage flight pitch angle, second-stage flight yaw angle, and second-stage flight roll angle of the aircraft after virtual ignition, wherein the second-stage flight pitch angle of the aircraft after virtual ignition is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle, and the second-stage flight yaw angle at the first iteration is the yaw angle at the optimal ignition time, and the second-stage flight roll angle is zero;

[0217] Calculate the first Z-axis error between the first Z-axis coordinate and the preset standard Z-axis coordinate;

[0218] The adjustment yaw angle is obtained by adding the second-stage flight yaw angle to the preset yaw angle adjustment step size. Based on the position and speed of the aircraft when the second-stage engine is virtually ignited at the optimal moment, the second-stage flight pitch angle, adjustment yaw angle and second-stage flight roll angle of the aircraft after virtual ignition are calculated. The second Z-axis coordinates of the aircraft when it is pushed out to the reentry point altitude are calculated.

[0219] Calculate the second Z-axis error between the second Z-axis coordinate and the standard Z-axis coordinate;

[0220] Based on the first Z-axis error and the second Z-axis error, calculate the derivative of the Z-axis error with respect to the yaw angle adjustment step size;

[0221] The secondary flight yaw angle is updated based on the yaw angle adjustment step size and the ratio of the Z-axis error to the derivative of the Z-axis error relative to the yaw angle adjustment step size. This process is iterated until the Z-axis error is less than the preset Z-axis error threshold or the preset number of iterations is reached. The secondary flight yaw angle after virtual ignition at the end of the iteration is taken as the optimal yaw angle.

[0222] Furthermore, in one embodiment, the second calculation module is also used to: cyclically calculate the first X-axis coordinates of the aircraft when it is extrapolated to the reentry point altitude based on the aircraft's position and velocity when the second-stage engine performs virtual ignition at the optimal ignition time, the second-stage pitch angle, the optimal yaw angle, and the second-stage roll angle of the aircraft after virtual ignition, wherein the second-stage pitch angle of the aircraft after virtual ignition is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle, and the final value of the pitch angle in the correspondence is the planned final value of the pitch angle, and the roll angle is zero;

[0223] Calculate the first X-axis error between the first X-axis coordinate and the preset standard X-axis coordinate;

[0224] The final pitch angle is obtained by adding the planned final pitch angle to the preset final pitch angle adjustment step size. Based on the position and velocity of the aircraft when the second-stage engine is virtually ignited at the optimal moment, the second-stage flight pitch angle, the optimal yaw angle and the second-stage flight roll angle of the aircraft after virtual ignition are calculated. The second X-axis coordinates of the aircraft when it is pushed out to the reentry point altitude are calculated.

[0225] Calculate the second X-axis error between the second X-axis coordinate and the standard X-axis coordinate;

[0226] Based on the first X-axis error and the second X-axis error, calculate the derivative of the adjustment step size of the X-axis error relative to the final value of the pitch angle;

[0227] The final pitch angle is updated based on the adjustment step size of the final pitch angle value and the ratio of the X-axis error to the derivative of the X-axis error relative to the final pitch angle value adjustment step size. This process is repeated until the X-axis error is less than the preset X-axis error threshold or the preset number of iterations is reached. The final pitch angle value after virtual ignition at the end of the iteration is taken as the optimal final pitch angle value.

[0228] Furthermore, in one embodiment, the second calculation module is also used to: determine the vector of the aircraft relative to the Earth's center after the second-stage engine of the aircraft has burned completely and the altitude of the aircraft is less than a preset altitude;

[0229] Calculate the dot product of the spacecraft's vector relative to the Earth's center and the spacecraft's velocity vector in the launch inertial frame;

[0230] Determine the magnitude of the vector of the aircraft relative to the Earth's center and the magnitude of the aircraft's velocity vector;

[0231] Based on the dot product, the magnitude of the aircraft's vector relative to the Earth's center, and the magnitude of the aircraft's velocity vector, calculate the reentry pitch angle, reentry yaw angle, and reentry roll angle when the reentry angle of attack satisfies the constrained angle of attack:

[0232]

[0233] γzr=0

[0234] in, ψzr is the reentry pitch angle, γzr is the reentry yaw angle, γzr is the reentry roll angle, Dxyzp is the dot product, Nxyz is the magnitude of the vector of the spacecraft relative to the Earth's center, and NVxyz is the magnitude of the spacecraft's velocity vector.

[0235] Furthermore, in one embodiment, the first computing module is further configured to:

[0236] When the location of a new target detection point is received, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the location of the new target detection point.

[0237] The functions of each module in the aforementioned energy planning device for the descent segment of an aircraft correspond to the steps in the aforementioned embodiment of the energy planning method for the descent segment of an aircraft, and their functions and implementation processes will not be described in detail here.

[0238] It should be noted that the sequence numbers of the embodiments in this application are for descriptive purposes only and do not represent the superiority or inferiority of the embodiments.

[0239] The terms "comprising" and "having," and any variations thereof, in the specification, claims, and accompanying drawings of this application are intended to cover non-exclusive inclusion. For example, a process, method, system, product, or apparatus that includes a series of steps or units is not limited to the listed steps or units, but may optionally include steps or units not listed, or may optionally include other steps or units inherent to such process, method, product, or apparatus. The terms "first," "second," and "third," etc., are used to distinguish different objects, etc., and do not indicate a sequence, nor do they limit "first," "second," and "third" to different types.

[0240] In the description of the embodiments of this application, terms such as "exemplary," "for example," or "for instance" are used to indicate examples, illustrations, or explanations. Any embodiment or design described as "exemplary," "for example," or "for instance" in the embodiments of this application should not be construed as being more preferred or advantageous than other embodiments or designs. Specifically, the use of terms such as "exemplary," "for example," or "for instance" is intended to present the relevant concepts in a concrete manner.

[0241] In the description of the embodiments of this application, unless otherwise stated, " / " means "or". For example, A / B can mean A or B. The "and / or" in the text is merely a description of the relationship between related objects, indicating that there can be three relationships. For example, A and / or B can mean: A exists alone, A and B exist simultaneously, and B exists alone. In addition, in the description of the embodiments of this application, "multiple" means two or more.

[0242] In some processes described in the embodiments of this application, multiple operations or steps are included in a specific order. However, it should be understood that these operations or steps may not be executed in the order they appear in the embodiments of this application, or they may be executed in parallel. The sequence number of the operation is only used to distinguish different operations, and the sequence number itself does not represent any execution order. In addition, these processes may include more or fewer operations, and these operations or steps may be executed sequentially or in parallel, and these operations or steps may be combined.

[0243] Through the above description of the embodiments, those skilled in the art can clearly understand that the methods of the above embodiments can be implemented by means of software plus necessary general-purpose hardware platforms. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on this understanding, the technical solution of this application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product is stored in a storage medium (such as ROM / RAM, magnetic disk, optical disk) as described above, and includes several instructions to cause a terminal device to execute the methods described in the various embodiments of this application.

[0244] The above are merely preferred embodiments of this application and do not limit the patent scope of this application. Any equivalent structural or procedural transformations made using the content of this application's specification and drawings, or direct or indirect applications in other related technical fields, are similarly included within the patent protection scope of this application.

Claims

1. A method for energy planning during the descent phase of an aircraft, characterized in that, The energy planning method for the descent segment of the aircraft includes: After the first-stage engine of the aircraft separates, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the position of the target detection point. When the aircraft's flight altitude exceeds the preset altitude, the aircraft's attitude is adjusted to the lowest RCS attitude, and the aircraft is controlled to fly without power. After the aircraft flies to the highest point of its orbit without power, the optimal firing time and optimal firing attitude of the second-stage engine are searched based on the orbital inclination and coordinate constraints of the reentry point. The system controls the second-stage engine of the aircraft to ignite at the optimal firing moment and adjusts the aircraft's attitude to the optimal firing attitude.

2. The energy planning method for the landing arc segment of an aircraft as described in claim 1, characterized in that, The calculation of the aircraft's minimum RCS attitude based on the aircraft's current position and the target detection point's position includes: Based on the current vector position of the aircraft in the launch inertial frame and the vector position of the target detection point in the launch inertial frame, determine the vector of the aircraft relative to the target detection point; The vector of the aircraft relative to the target detection point is vector normalized to obtain the corresponding first unit vector; The current vector position of the aircraft is vector normalized to obtain the corresponding second unit vector; The reference axis is obtained by performing the cross product of the first unit vector and the second unit vector in the opposite direction. The cross product of the reference axis and the first unit vector is calculated to obtain the corrected reference axis. Based on the first unit vector, the corrected reference axis, and the reference axis, calculate the minimum RCS pitch angle, minimum RCS yaw angle, and minimum RCS roll angle in the minimum RCS attitude: ψ = a sin(-Zx) γ = a tan2(Zy, Zz) in, ψ is the lowest RCS pitch angle, γ is the lowest RCS yaw angle, γ is the lowest RCS roll angle, Xx is the x-component of the first unit vector, Yx is the x-component of the corrected reference axis, Zx is the x-component of the reference axis, Zy is the y-component of the reference axis, and Zz is the z-component of the reference axis.

3. The energy planning method for the landing arc segment of an aircraft as described in claim 1, characterized in that, The process of searching for the optimal firing time and optimal firing attitude of the second-stage engine of the aircraft, based on the orbital inclination angle constraint and reentry point coordinate constraint of the reentry point, includes: Based on the orbital inclination constraints of the spacecraft, search for the optimal firing time of the spacecraft's second-stage engine; Based on the Z-axis coordinate constraint of the aircraft's reentry point, search for the optimal yaw angle of the aircraft after the second-stage engine ignition; Based on the X-axis coordinate constraint of the aircraft's reentry point, search for the final value of the aircraft's optimal pitch angle after the second-stage engine ignites.

4. The energy planning method for the landing arc segment of an aircraft as described in claim 3, characterized in that, Before searching for the optimal firing time and optimal firing attitude of the second-stage engine based on the orbital inclination and coordinate constraints of the reentry point, the following steps are also included: After the second-stage engine of the aircraft is set to ignite, the correspondence between the combustion process of the second-stage engine and the pitch angle of the aircraft is as follows: For the first third of the second-stage engine's combustion period, the aircraft's pitch angle remains the initial pitch angle at the time of the second-stage engine's ignition. During the middle third of the second-stage engine combustion period, the aircraft pitch angle is linearly adjusted from the initial pitch angle to the final pitch angle value. During the last third of the second-stage engine's combustion period, the aircraft maintains its final pitch angle value.

5. The energy planning method for the landing arc segment of an aircraft as described in claim 4, characterized in that, The search for the optimal firing moment of the second-stage engine of the spacecraft, based on the spacecraft's orbital inclination constraints, includes: The system calculates the first inclination angle of the orbit when the aircraft is pushed outward to the reentry point altitude based on the aircraft's position and velocity when the second-stage engine is virtually ignited at the planned ignition time, the second-stage flight pitch angle, the second-stage flight yaw angle, and the second-stage flight roll angle after the virtual ignition. The second-stage flight pitch angle is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle, the second-stage flight yaw angle is the yaw angle at the virtual ignition time, and the second-stage flight roll angle is zero. Calculate the inclination error between the first inclination angle of the track and the preset standard inclination angle value; The adjusted ignition time is obtained by adding the preset time adjustment step to the planned ignition time. Based on the position and speed of the aircraft when the second-stage engine is virtually ignited at the adjusted ignition time, and the second-stage flight pitch angle, second-stage flight yaw angle and second-stage flight roll angle of the aircraft after virtual ignition, the second inclination angle of the trajectory when the aircraft is pushed out to the reentry point altitude is calculated. Calculate the derivative of the orbit inclination angle relative to the time adjustment step based on the first and second orbit inclination angles; The planned ignition time is updated based on the time adjustment step size and the ratio of the inclination error to the derivative of the orbital inclination with respect to the time adjustment step size. This process is iterated until the inclination error is less than the preset inclination error threshold or the preset number of iterations is reached. The virtual ignition time corresponding to the end of the iteration is taken as the optimal ignition time.

6. The energy planning method for the landing arc segment of an aircraft as described in claim 4, characterized in that, The process of searching for the optimal yaw angle of the aircraft after the second-stage engine ignition, based on the Z-axis coordinate constraints of the aircraft's reentry point, includes: The loop calculates the first Z-axis coordinates of the aircraft when it is pushed to the reentry point altitude based on the aircraft's position and velocity when the second-stage engine is virtually ignited at the optimal ignition time, the second-stage flight pitch angle, second-stage flight yaw angle and second-stage flight roll angle after the virtual ignition. The second-stage flight pitch angle of the aircraft after the virtual ignition is obtained by time interpolation based on the correspondence between the second-stage engine combustion process and the pitch angle. The second-stage flight yaw angle at the first iteration is the yaw angle at the optimal ignition time, and the second-stage flight roll angle is zero. Calculate the first Z-axis error between the first Z-axis coordinate and the preset standard Z-axis coordinate; The adjustment yaw angle is obtained by adding the second-stage flight yaw angle to the preset yaw angle adjustment step size. Based on the position and speed of the aircraft when the second-stage engine is virtually ignited at the optimal moment, the second-stage flight pitch angle, adjustment yaw angle and second-stage flight roll angle of the aircraft after virtual ignition are calculated. The second Z-axis coordinates of the aircraft when it is pushed out to the reentry point altitude are calculated. Calculate the second Z-axis error between the second Z-axis coordinate and the standard Z-axis coordinate; Based on the first Z-axis error and the second Z-axis error, calculate the derivative of the Z-axis error with respect to the yaw angle adjustment step size; The secondary flight yaw angle is updated based on the yaw angle adjustment step size and the ratio of the Z-axis error to the derivative of the Z-axis error relative to the yaw angle adjustment step size. This process is iterated until the Z-axis error is less than the preset Z-axis error threshold or the preset number of iterations is reached. The secondary flight yaw angle after virtual ignition at the end of the iteration is taken as the optimal yaw angle.

7. The energy planning method for the landing arc segment of an aircraft as described in claim 4, characterized in that, The process of searching for the final value of the optimal pitch angle of the aircraft after the second-stage engine ignition, based on the X-axis coordinate constraints of the aircraft's reentry point, includes: The loop calculates the first X-axis coordinates of the aircraft when it is pushed out to the reentry point altitude based on the aircraft's position and velocity when the second-stage engine is ignited at the optimal moment, the second-stage pitch angle, the optimal yaw angle, and the second-stage roll angle after the virtual ignition. The second-stage pitch angle of the aircraft after the virtual ignition is obtained by interpolation over time based on the correspondence between the second-stage engine combustion process and the pitch angle. In the correspondence, the final pitch angle is the planned final pitch angle, and the roll angle is zero. Calculate the first X-axis error between the first X-axis coordinate and the preset standard X-axis coordinate; The final pitch angle is obtained by adding the planned final pitch angle to the preset final pitch angle adjustment step size. Based on the position and velocity of the aircraft when the second-stage engine is virtually ignited at the optimal moment, the second-stage flight pitch angle, the optimal yaw angle and the second-stage flight roll angle of the aircraft after virtual ignition are calculated. The second X-axis coordinates of the aircraft when it is pushed out to the reentry point altitude are calculated. Calculate the second X-axis error between the second X-axis coordinate and the standard X-axis coordinate; Based on the first X-axis error and the second X-axis error, calculate the derivative of the adjustment step size of the X-axis error relative to the final value of the pitch angle; The final pitch angle is updated based on the adjustment step size of the final pitch angle value and the ratio of the X-axis error to the derivative of the X-axis error relative to the final pitch angle value adjustment step size. This process is repeated until the X-axis error is less than the preset X-axis error threshold or the preset number of iterations is reached. The final pitch angle value after virtual ignition at the end of the iteration is taken as the optimal final pitch angle value.

8. The energy planning method for the landing arc segment of an aircraft as described in claim 1, characterized in that, The method also includes: After the second-stage engine of the aircraft has burned completely and the aircraft's altitude is less than the preset altitude, determine the vector of the aircraft relative to the Earth's center; Calculate the dot product of the spacecraft's vector relative to the Earth's center and the spacecraft's velocity vector in the launch inertial frame; Determine the magnitude of the vector of the aircraft relative to the Earth's center and the magnitude of the aircraft's velocity vector; Based on the dot product, the magnitude of the aircraft's vector relative to the Earth's center, and the magnitude of the aircraft's velocity vector, calculate the reentry pitch angle, reentry yaw angle, and reentry roll angle when the reentry angle of attack satisfies the constrained angle of attack: γzr=0 in, ψzr is the reentry pitch angle, γzr is the reentry yaw angle, γzr is the reentry roll angle, Dxyzp is the dot product, Nxyz is the magnitude of the vector of the spacecraft relative to the Earth's center, and NVxyz is the magnitude of the spacecraft's velocity vector.

9. The energy planning method for the landing arc segment of an aircraft as described in claim 1, characterized in that, After the first-stage engine of the aircraft separates, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the position of the target detection point, including: When the location of a new target detection point is received, the minimum RCS attitude of the aircraft is calculated based on the current position of the aircraft and the location of the new target detection point.

10. An energy planning device for the landing arc segment of an aircraft, characterized in that, The energy planning device for the landing arc segment of the aircraft includes: The first calculation module is used to calculate the minimum RCS attitude of the aircraft based on the current position of the aircraft and the position of the target detection point after the separation of the first stage engine of the aircraft. The first adjustment module is used to adjust the attitude of the aircraft to the lowest RCS attitude and control the aircraft to fly without power when the aircraft's flight altitude is greater than the preset altitude. The second calculation module is used to search for the optimal firing time and optimal firing attitude of the second-stage engine of the aircraft after the aircraft flies to the highest point of the orbit without power, based on the orbital inclination angle constraint and the reentry point coordinate constraint of the aircraft. The second adjustment module is used to control the second-stage engine of the aircraft to ignite at the optimal firing moment and adjust the attitude of the aircraft to the optimal firing attitude.

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