Aircraft arc lowering section energy planning method and device

Abstract

The invention provides an aircraft arc lowering section energy planning method and device, and the method comprises the steps: calculating the lowest RCS attitude of an aircraft according to the current position of the aircraft and the position of a target detection point after the separation of a first-stage engine of the aircraft; when the flight height of the aircraft is greater than a preset height, adjusting the attitude of the aircraft to the lowest RCS attitude, and controlling the aircraft to perform unpowered flight; after the aircraft flies to the highest point of the orbit in an unpowered mode, the optimal ignition moment and the optimal ignition attitude of a secondary engine of the aircraft are searched according to the inclination angle constraint and the reentry point coordinate constraint of the aircraft; the secondary engine of the aircraft is controlled to ignite at the optimal ignition moment, and the attitude of the aircraft is adjusted to the optimal ignition attitude, so that the secondary engine of the aircraft is controlled not to ignite at the rising arc section of the elliptical orbit; the aircraft flies according to the searched optimal ignition moment and the optimal ignition attitude, and flies according to the attitude constrained by the reentry attack angle after the engine is exhausted, so that the aircraft orbit is ensured to meet the constraint, and the difficulty of detection tracking and orbit prediction of the aircraft in the secondary flight section is improved.

Description

Technical Field

[0001] The present application relates to the technical field of aircraft guidance and control, and in particular to a method and device for energy planning of an aircraft descent arc segment. Background Art

[0002] The re-entry glide vehicle combines the orbital characteristics of the re-entry vehicle and the orbital characteristics of the gliding vehicle. After the first-stage flight segment, it can use the second-stage solid engine to manage energy, allowing the vehicle to break through the atmosphere in a short time, fly to the re-entry point in outer space in a re-entry orbit, and then pull up and glide after entering the re-entry point, thereby achieving long-distance flight and high-precision landing.

[0003] However, there are still some problems with the two-stage guidance design of re-entry glide vehicles in the related art. Traditional two-stage guidance designs mainly use closed-loop guidance methods, igniting the engine during the ascent phase of the elliptical orbit and guiding and controlling the engine energy so that the re-entry point of the aircraft meets the orbital inclination, position error, angle of attack, etc. However, although this design can achieve precise control of the aircraft's re-entry point to a certain extent, it also has obvious drawbacks. When the engine is ignited during the ascent phase of the elliptical orbit, the aircraft will produce obvious infrared and radar signatures, which makes the aircraft easily observable by detection equipment, thereby exposing the aircraft's position and flight trajectory, and the aircraft's re-entry point is easy to predict.

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

[0005] The present application provides a method and device for energy planning of the descent arc segment of an aircraft, which can solve the technical problem in the prior art that when an aircraft ignites its engine in the ascending segment of an elliptical orbit, the subsequent flight trajectory can be easily detected and tracked.

[0006] In a first aspect, an embodiment of the present application provides an energy planning method for an aircraft descending arc segment, the energy planning method for an aircraft descending arc segment comprising:

[0007] After the first stage engine of the aircraft separates, the lowest 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 aircraft flies without power to the highest point of the orbit, according to the orbit inclination constraint and the reentry point coordinate constraint of the aircraft, search for the optimal ignition time and the optimal ignition attitude of the secondary engine of the aircraft;

[0010] Control the secondary engine of the aircraft to ignite at the optimal ignition time, and adjust the attitude of the aircraft to the optimal ignition attitude.

[0011] In a second aspect, an embodiment of the present application provides an energy planning device for the descending arc section of an aircraft. The energy planning device for the descending arc section of the aircraft includes:

[0012] A first calculation module, which is used to calculate the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the target detection point after the separation of the primary engine of the aircraft;

[0013] A first adjustment module, which is used to adjust the attitude of the aircraft to the lowest RCS attitude when the flight altitude of the aircraft is greater than a preset altitude, and control the aircraft to fly without power;

[0014] A second calculation module, which is used to search for the optimal ignition time and the optimal ignition attitude of the secondary engine of the aircraft according to the orbit inclination constraint and the reentry point coordinate constraint of the aircraft after the aircraft flies without power to the highest point of the orbit;

[0015] A second adjustment module, which is used to control the secondary engine of the aircraft to ignite at the optimal ignition time, and adjust the attitude of the aircraft to the optimal ignition attitude.

[0016] The beneficial effects brought by the technical solution provided by the embodiment of the present application include:

[0017] After the separation of the primary engine of the aircraft, calculate the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the target detection point; when the flight altitude of the aircraft is greater than the preset altitude, adjust the attitude of the aircraft to the lowest RCS attitude, and control the aircraft to fly without power; after the aircraft flies without power to the highest point of the orbit, search for the optimal ignition time and the optimal ignition attitude of the secondary engine of the aircraft according to the orbit inclination constraint of the reentry point and the reentry point coordinate constraint; control the secondary engine of the aircraft to ignite at the optimal ignition time, and adjust the attitude of the aircraft to the optimal ignition attitude. It is realized that after the primary engine of the aircraft burns out and separates, the secondary engine does not ignite in the ascending arc section of the elliptical orbit, but rises to a certain height and then adjusts to the lowest RCS attitude for the detection point. After flying over the highest point of the elliptical orbit without power, fly according to the searched optimal ignition time and the optimal ignition attitude, and fly according to the attitude constrained by the reentry attack angle after the engine runs out, ensuring that the aircraft orbit meets the constraints, thereby increasing the difficulty of detecting, tracking and orbit prediction of the aircraft in the secondary flight section. Brief Description of the Drawings

[0018] Figure 1 It is a schematic flowchart of an embodiment of the energy planning method for the descending arc section of the aircraft in this application;

[0019] Figure 2 It is a schematic diagram of the stage division of the aircraft re - entry glide orbit;

[0020] Figure 3 It is a schematic diagram of the corresponding relationship between the combustion process of the secondary engine and the pitch angle of the aircraft;

[0021] Figure 4 It is a schematic diagram of the functional modules of an embodiment of the energy planning device for the descending arc section of the aircraft in this application. Detailed Embodiment

[0022] In order to enable those skilled in the art to better understand the solution of this application, the technical solutions in the embodiments of this application will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of this application. Obviously, the described embodiments are only a part of the embodiments of this application, rather than all of the embodiments. Based on the embodiments in this application, all other embodiments obtained by those of ordinary skill in the art without making creative efforts belong to the scope of protection of this application.

[0023] First, some technical terms in this application are explained to facilitate the understanding of this application by those skilled in the art.

[0024] Lowest RCS attitude: It refers to an attitude of the aircraft during flight, in which the aircraft adjusts its own attitude to make its Radar Cross Section (RCS) the smallest.

[0025] To make the purpose, technical solution and advantages of this application clearer, the following will further describe the embodiments of this application in detail in conjunction with the accompanying drawings.

[0026] In the first aspect, the embodiments of this application provide an energy planning method for the descending arc section of an aircraft.

[0027] In one embodiment, please refer to Figure 1 and Figure 2 , Figure 1 is a schematic flowchart of the first embodiment of the energy planning method for the descending arc section of the aircraft in this application. As Figure 1 shown, the energy planning method for the descending arc section of the aircraft includes:

[0028] Step S1: After the separation of the primary engine of the aircraft, calculate the lowest RCS attitude of the aircraft according to 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 vector of the aircraft relative to the target detection point based on the current vector position of the aircraft in the launch inertial system and the vector position of the target detection point in the launch inertial system :

[0031]

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

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

[0034] wherein, the modulus of the vector of the aircraft relative to the target detection point is:

[0035]

[0036] wherein, is the modulus of, is the component in the X direction, is the component in the Y direction, is the component in the Z direction.

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

[0038]

[0039] wherein, is the first unit vector.

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

[0041] wherein, the modulus of the current vector position of the aircraft is:

[0042]

[0043] wherein, is the modulus of, is the component in the X direction, is the component in the Y direction, is The component in the Z direction.

[0044] For Perform vector normalization to obtain the corresponding second unit vector:

[0045]

[0046] Wherein, Is the second unit vector.

[0047] Step S104: Perform a cross product calculation on the first unit vector and the second unit vector in the opposite direction to obtain the reference axis:

[0048]

[0049] Wherein, Is 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] Wherein, Is the corrected reference axis.

[0053] Step S106: Calculate the lowest RCS pitch angle, the lowest RCS yaw angle, and the lowest RCS roll angle in the lowest RCS attitude according to the first unit vector, the corrected reference axis, and the reference axis:

[0054]

[0055] ψ = a sin(-Zx)

[0056] γ = a tan 2(Zy, Zz)

[0057] Wherein, Is the lowest RCS pitch angle, ψ 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 embodiment, an A matrix can be established according to the first unit vector, the corrected reference axis, and the reference axis:

[0059]

[0060] Calculate the lowest RCS pitch angle, the lowest RCS yaw angle, and the lowest RCS roll angle according to the A matrix:

[0061]

[0062] ψ = a sin(-A13)

[0063] γ = a tan 2(A23, A33)

[0064] Wherein, 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 separation of the first-stage engine of the aircraft, according to the current position of the aircraft and the position of the target detection point, calculate the minimum RCS attitude of the aircraft, including: when receiving the position of a new target detection point, calculate the minimum RCS attitude of the aircraft according to the current position of the aircraft and the position of the new target detection point.

[0066] Explanatorily, the detection point in this embodiment may be a radar detection point. During the flight of the aircraft, multiple detection points may be encountered. One of the multiple detection points can be selected as the target detection point, and during the flight of the aircraft, it is supported to update the target detection point and its position to calculate the minimum RCS attitude of the aircraft relative to the target detection point in real time, thereby increasing the difficulty of detecting and tracking the aircraft.

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

[0068] It should be noted that the preset altitude in this embodiment can be set to 70KM or 100KM. By setting that after the first-stage active section engine of the aircraft burns out and separates, the second-stage engine does not ignite during the ascending arc section of the elliptical orbit but rises to the preset altitude and then adjusts to the minimum RCS attitude for the detection point, the concealment of the aircraft can be improved and the difficulty of its detection and tracking can be increased.

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

[0070] In one embodiment, before searching for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbital inclination constraint and reentry point coordinate constraint of the reentry point of the aircraft, it further includes:

[0071] Such as Figure 3As shown, after setting the ignition of the second-stage engine of the aircraft, the corresponding relationship between the combustion process of the second-stage engine and the pitch angle of the aircraft is as follows: The time between the ignition of the second-stage engine and the completion of the combustion of the second-stage engine is the engine combustion time (Tfire). The second-stage engine combustion time (Tfire) is divided into three segments. During the first one-third of the time of the second-stage engine combustion of the aircraft, that is, 0 to 0.333*Tfire, it is the attitude stabilization stage of the aircraft, and the pitch angle of the aircraft remains the initial pitch angle at the ignition of the second-stage engine. During the middle one-third of the time of the second-stage engine combustion of the aircraft, that is, 0.333*Tfire to 0.666*Tfire, it is the alternating attitude section of the aircraft, and the pitch angle of the aircraft is linearly adjusted from the initial pitch angle to the final pitch angle. During the last one-third of the time of the second-stage engine combustion of the aircraft, that is, 0.666*Tfire to Tfire, it is the attitude holding section of the aircraft, and the pitch angle of the aircraft remains the final pitch angle. Therefore, as Figure 3 shown, according to the corresponding relationship between the combustion process of the second-stage engine and the pitch angle of the aircraft, the corresponding pitch angle of the aircraft can be obtained by time interpolation.

[0072] Specifically, step S3 searches for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbital inclination constraint and the re-entry point coordinate constraint of the aircraft, and specifically includes the following steps: According to the orbital inclination constraint of the aircraft, search for the optimal ignition time of the second-stage engine of the aircraft; according to the Z-direction coordinate constraint of the re-entry point of the aircraft, search for the optimal yaw angle of the aircraft after the ignition of the second-stage engine; according to the X-direction coordinate constraint of the re-entry point of the aircraft, search for the optimal final pitch angle of the aircraft after the ignition of the second-stage engine.

[0073] In one embodiment, searching for the optimal ignition time of the second-stage engine of the aircraft according to the orbital inclination constraint of the aircraft includes:

[0074] Step S311: Loop starting from the current time T, the aircraft performs unpowered extrapolation to the planned ignition time Tign for virtual ignition. According to the position and velocity of the aircraft during virtual ignition at the planned ignition time of the second-stage engine, 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. Then, the aircraft performs powered extrapolation until the second-stage engine burns out and then switches to unpowered extrapolation until the corresponding first orbital inclination θzr1 at the re-entry point height Hend is reached.

[0075] Among them, the second-stage flight pitch angle is obtained by time interpolation according to the corresponding relationship between the second-stage engine combustion process and the pitch angle. At this time, in the corresponding relationship between the second-stage engine combustion process and the pitch angle, the final value of the pitch angle is the planned final value of the pitch angle, that is, the final value of the pitch angle designed in the standard orbit of the aircraft. 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. At the first iteration, the planned ignition moment Tign is the designed second-stage engine ignition moment in the standard orbit of the aircraft.

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

[0077] dθzr = θzr1 - θbz

[0078] Among them, dθzr is the inclination error, θzr1 is the first orbit inclination angle, and θbz is the standard inclination value. The standard inclination value is the orbit inclination angle of the reentry point designed in the standard orbit of the aircraft.

[0079] Step S313: Add the preset time adjustment step dT to the planned ignition moment Tign to obtain the adjusted ignition moment Tign + dT. Starting from the current moment T, the aircraft performs unpowered extrapolation to the adjusted ignition moment Tign + dT for virtual ignition. According to the position and velocity of the aircraft at the adjusted ignition moment during virtual ignition of the second-stage engine, the second-stage flight pitch angle, second-stage flight yaw angle, and second-stage flight roll angle of the aircraft after virtual ignition, calculate the aircraft to perform powered extrapolation until the second-stage engine burns out and then switches to unpowered extrapolation until the corresponding second orbit inclination angle θzr2 at the reentry point height Hend is reached.

[0080] The obtaining methods of the second-stage flight pitch angle, second-stage flight yaw angle, and second-stage flight roll angle are the same as those in step S311 and will not be elaborated here.

[0081] Step S314: Calculate the derivative of the orbit inclination angle with respect to the time adjustment step according to the first orbit inclination angle and the second orbit inclination angle:

[0082] dθzrdt = (θzr2 - θzr11) / dT

[0083] Among them, θzr1 is the first orbit inclination angle, θzr2 is the second orbit inclination angle, 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: Update the planned ignition moment according to the time adjustment step and the ratio of the inclination error to the derivative of the orbit inclination angle with respect to the time adjustment step:

[0085] Tign = Tign - dT - dθzr / dθzrdt

[0086] After the updated planned ignition time, steps S311 to S315 are continuously looped for iteration until the inclination error is less than the preset inclination error threshold or the preset number of iterations is reached, and the corresponding virtual ignition time at the end of the iteration is used as the optimal ignition time.

[0087] Exemplarily, the inclination error threshold in this embodiment is 0.1°, and the preset number of iterations is 10 times, that is, until the number of loops is greater than 10 times or it is calculated that dθzr < 0.1°. The virtual ignition time obtained at this time is the optimal ignition time Tign that satisfies the reentry point orbit inclination constraint. opt 。

[0088] In one embodiment, according to the Z - coordinate constraint of the reentry point of the aircraft, searching for the optimal yaw angle of the aircraft after the second - stage engine ignition includes:

[0089] Step S321: Loop starting from the current time T, the aircraft performs unpowered extrapolation to the optimal ignition time Tign. opt Perform virtual ignition. According to the position and velocity of the aircraft when virtual ignition is performed by the second - stage engine at the optimal ignition time, the corresponding second - stage flight pitch angle, second - stage flight yaw angle, and second - stage flight roll angle of the aircraft after virtual ignition, calculate the aircraft's powered extrapolation until the second - stage engine burns out and then switches to unpowered extrapolation until the first Z - coordinate Z1 corresponding to the reentry point height Hend is reached.

[0090] Among them, the second - stage flight pitch angle of the aircraft after virtual ignition is obtained by time interpolation according to the corresponding relationship between the second - stage engine combustion process and the pitch angle. In the corresponding relationship between the second - stage engine combustion process and the pitch angle at this time, the final value of the pitch angle is the planned final value of the pitch angle, that is, the final value of the pitch angle designed in the standard orbit of the aircraft. The second - stage flight yaw angle at the first iteration maintains the yaw angle ψ at the optimal ignition time, and the second - stage flight roll angle is zero.

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

[0092] dZ1 = Z1 - Zbz

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

[0094] Step S323: Add the preset yaw angle adjustment step dψ to the secondary flight yaw angle ψ to obtain the adjusted yaw angle ψ + dψ. According to the position and velocity of the aircraft during virtual ignition at the optimal ignition time of the secondary engine, the secondary flight pitch angle, the adjusted yaw angle ψ + dψ, and the secondary flight roll angle corresponding to the aircraft after virtual ignition, calculate the aircraft with power extrapolation until the secondary engine burns out and then switches to powerless extrapolation until the second Z - coordinate Z2 corresponding to the re - entry point altitude Hend is reached.

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

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

[0097] dZ2 = Z2 - Zbz

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

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

[0100] dZdψ=(dZ2 - dZ1) / dψ

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

[0102] Step S326: According to the yaw angle adjustment step and the ratio of the Z - direction error to the derivative of the Z - direction error with respect to the yaw angle adjustment step, update the planned yaw angle:

[0103] ψ = ψ - dψ - dZ / dZdψ

[0104] After updating the planned yaw angle, continue to loop from step S321 to step S326 for iteration until the Z - direction error is less than the preset Z - direction error threshold or the preset iteration times are reached, and use the yaw angle after virtual ignition corresponding to the end of the iteration as the optimal yaw angle.

[0105] Exemplarily, in this embodiment, the Z - direction error threshold is 500 meters and the preset iteration times are 10 times, that is, until the number of loops is greater than 10 times or dZ < 500 meters. At this time, the optimal yaw angle ψ that satisfies the Z - direction error constraint after secondary ignition is obtained. opt 。

[0106] In one embodiment, according to the reentry point X-direction coordinate constraint of the aircraft, searching for the optimal pitch angle final value of the aircraft after the second-stage engine ignition includes:

[0107] Step S331: Loop starting from the current time T, the aircraft performs unpowered extrapolation to the optimal ignition time Tign opt Perform virtual ignition. According to the position and velocity of the aircraft during virtual ignition at the optimal ignition time of the second-stage engine, the corresponding second-stage flight pitch angle ψ of the aircraft after virtual ignition opt , the optimal yaw angle and roll angle ψ opt , calculate that the aircraft performs powered extrapolation until the second-stage engine burns out and then switches to unpowered extrapolation until the corresponding first X-direction coordinate X1 at the reentry point altitude Hend is reached.

[0108] Among them, the second-stage flight pitch angle of the aircraft after virtual ignition is obtained by time interpolation according to the corresponding relationship between the second-stage engine combustion process and the pitch angle. In the first iteration, the final pitch angle value in the corresponding relationship between the second-stage engine combustion process and the pitch angle is the planned final pitch angle value, that is, the final pitch angle value designed in the standard orbit of the aircraft. The second-stage flight roll angle is zero.

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

[0110] dX1 = X1 - Xbz

[0111] Among them, dX1 is the first X-direction error, X1 is the first X-direction coordinate, and Xbz is the standard X-direction coordinate. The standard X-direction coordinate is the reentry point X-direction coordinate designed in the standard orbit of the aircraft.

[0112] Step S333: Add the planned final pitch angle value to the preset final pitch angle value adjustment step size to obtain the adjusted final pitch angle value According to the position and velocity of the aircraft during virtual ignition at the optimal ignition time of the second-stage engine, the corresponding second-stage flight pitch angle, optimal yaw angle ψ opt and the second-stage flight roll angle of the aircraft after virtual ignition, calculate that the aircraft performs powered extrapolation until the second-stage engine burns out and then switches to unpowered extrapolation until the corresponding second X-direction coordinate X2 at the reentry point altitude Hend is reached.

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

[0114] dX2 = X2 - Xbz

[0115] Among them, dX2 is the second X error.

[0116] Step S335: Calculate the derivative of the X-direction error relative to the pitch angle final value adjustment step length based on the first X-direction error and the second X-direction error:

[0117]

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

[0119] Step S336: Update the planned final pitch angle value according to the pitch angle final value adjustment step size and the ratio of the X-direction error to the derivative of the X-direction error with respect to the pitch angle final value adjustment step size:

[0120]

[0121] After the planned final pitch angle value is updated, the process continues with steps S331 to S336 and iterates until the X-axis error is less than a preset X-axis error threshold or the preset number of iterations is reached, and the final yaw angle value corresponding to the virtual ignition at the end of the iteration is used as the optimal final pitch angle value.

[0122] For example, in this embodiment, the X-axis error threshold is 500 meters, and the number of iterations is preset to 10, that is, until the number of cycles is greater than 10 or dX is calculated to be less than 500 meters. The optimal final pitch angle after the second stage ignition that meets the X-axis error constraint at the reentry point is obtained.

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

[0124] Step S341: Calculate the current aircraft gravitational acceleration Gxyz k , initial time k=0,t k =0;

[0125] Step S342, set the extrapolation step length step = 0.02s;

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

[0127]

[0128] Step S344: Calculate the acceleration Gxyz of the current position based on the updated position result. k+1 , calculate the velocity 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 unpowered gliding to Tign.

[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 unpowered gliding to the target height Tend.

[0132] In one embodiment, after the second-stage engine of the aircraft ignites, it is extrapolated with power until the completion of the second-stage engine combustion. 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 k at the current moment and the apparent acceleration a k+1 at the next moment. Then there is:

[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, according to the updated position results, calculate the acceleration Gxyz at the current position k+1 , and 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 , end extrapolation to obtain the flight time T at the moment when the second-stage engine runs out endwt = t k + T0 + T start 、Position velocity XYZ endwt Vxyz endwt .

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

[0145] As a preferred implementation, the method further includes:

[0146] [[ID=3I]]After the second-stage engine of the aircraft burns out completely and the altitude of the aircraft is less than the preset altitude, adjust the attitude to the attitude with the reentry attack angle being the constrained attack angle α0 according to the current position velocity, and its calculation steps are as follows:

[0147] Determine the vector of the aircraft relative to the earth's center:

[0148]

[0149] Wherein, is the vector of the aircraft relative to the earth's center, and Rox, Roy, and Roz are the three components of the earth's center in the launch inertial system respectively.

[0150] Calculate the dot product of the vector of the aircraft relative to the earth's center and the velocity vector of the aircraft in the launch inertial system ]>:

[0151] Dxyzp = XpVx + YpVy + ZpVz

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

[0153] Determine the modulus of the vector of the aircraft relative to the earth's center and the modulus of the velocity vector of the aircraft:

[0154]

[0155] Wherein, Nxyz is the modulus of the vector of the aircraft relative to the earth's center, and NVxyz is the modulus of the velocity vector of the aircraft.

[0156] According to the dot product, the magnitude of the vector of the aircraft relative to the center of the earth, and the magnitude of the velocity vector of the aircraft, calculate the reentry pitch angle, reentry yaw angle, and reentry roll angle when the reentry attack angle of the aircraft satisfies the constraint attack angle:

[0157]

[0158] γzr = 0

[0159] Wherein, is the reentry pitch angle, ψzr 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 aircraft relative to the center of the earth, and NVxyz is the magnitude of the velocity vector of the aircraft.

[0160] It should be noted that after the aircraft is controlled to pass through the highest point of the unpowered flyover orbit, the method of virtual ignition and open-loop prediction is adopted to plan the appropriate ignition timing and energy dissipation attitude to ensure the orbit inclination and reentry point error constraints. This method can effectively increase the difficulty of detecting, tracking, and orbit prediction of the reentry-glide aircraft in the second-stage flight section and meet the requirement constraints to enter the glide section.

[0161] In a specific embodiment, assume that the current position of an aircraft is (629312; 148365; 11404 m); the position of the target radar observation station is (1064700; -60478; -90723 m); 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 entering the altitude of 70 Km, according to the current position of the aircraft and the position of the radar station, calculate the lowest RCS attitude. First, calculate the reference axis Xx = (0.88; -0.42 -0.21); calculate Yref = (0.97; 0.23 0.02), calculate Zz = (-0.04; 0.22; -0.61), calculate the corrected reference Y axis Yy = (-0.30; -0.55; -0.17); form matrix A with Xx, Yy, and Zz, and solve for the lowest RCS attitude to be a pitch angle of -19.06°; a yaw angle of 2.29°, and a roll angle of 160.54°. In this attitude, the aircraft will face the radar station head-on, making the detection RCS value of the aircraft for the radar station the lowest.

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

[0164] The re-entry point coordinate position in the standard orbit design is (914951; -3062; -39372), the required orbit inclination angle at the re-entry point is -15.8° ± 1°, and the re-entry is maintained at a zero angle of attack. Energy planning for the descending arc section is carried out for this.

[0165] First, as Figure 3 shown, the combustion time of the second-stage engine is divided into three equal parts, which are the attitude stabilization section in the early stage of ignition, the alternating attitude section in the middle stage of ignition, and the attitude holding section in the late stage of ignition; after ignition, interpolation is performed according to the engine combustion time to obtain the corresponding pitch angle attitude.

[0166] Then, the ignition time is optimized to obtain the optimal ignition time, and a loop is constructed. According to the current position and speed, extrapolate without power to the virtual ignition at 388.012 s, and then extrapolate with power. During this period, the pitch angle of the second-stage flight is obtained by interpolation according to 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 for power extrapolation; when the second-stage engine burns out, continue to extrapolate without power until the Y-axis is less than or equal to the Y-axis of the re-entry point (-3062 m) and then stop, and calculate that the orbit inclination angle at the re-entry point is -16.104°.

[0167] Then, the time adjustment step size is designed to be 0.05 s. According to the current position and speed, extrapolate without power to the ignition at 388.062 s; then extrapolate with power. During this period, the pitch angle of the second-stage flight is obtained by interpolation according to 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 for power extrapolation; when the second-stage engine burns out, continue to extrapolate without power until the Y-axis is less than or equal to the Y-axis of the re-entry point (-3062 m) and then stop, and calculate that the orbit inclination angle at the re-entry point is -16.098°.

[0168] So far, it is calculated that dθzrdt = 0.1122; the ignition time for the next cycle is 390.653 s;

[0169] When the number of iterations > 10 times, or the deviation of the re-entry inclination angle from the standard orbit constraint < 0.1°, jump out of the loop to obtain the optimal ignition time of 390.653 s; the number of loops is 1 time.

[0170] Furthermore, optimize the yaw angle to obtain the optimal yaw angle and construct a loop. Extrapolate without power to the optimal ignition time of 390.653 s for ignition according to the current position speed. Then extrapolate with power. During this period, the pitch angle of the second-stage flight is obtained by interpolation according to the burning time of the second-stage engine after ignition. The yaw angle of the second-stage flight is 30.8°, and the roll angle of the second-stage flight is 0. Then perform extrapolation with power. When the second-stage engine burns out, continue to extrapolate without power until the Y-axis is less than or equal to the Y-axis of the reentry point (-3062 m) and stop. Calculate that the Z-direction position deviation of the reentry point is 4296 m.

[0171] Then design the yaw angle adjustment step size to be 0.1°. Extrapolate without power to the ignition at 390.653 s according to the current position speed. Then extrapolate with power. During this period, the pitch angle of the second-stage flight is obtained by interpolation according to the burning time of the second-stage engine after ignition. Adjust the yaw angle to 30.9°, and the roll angle of the second-stage flight is 0. Then perform extrapolation with power. When the second-stage engine burns out, continue to extrapolate without power until the Y-axis is less than or equal to the Y-axis of the reentry point (-3062 m) and stop. Calculate that the Z-direction position deviation of the reentry point is 4153 m.

[0172] Thus, it is calculated that dZdψ = -81837.5; the yaw angle of the second-stage flight in the next loop is 32.907°;

[0173] When the number of iterations > 10 times, or the Z-direction reentry point deviation from the standard orbit constraint < 500 m, jump out of the loop to obtain the optimal yaw angle of 32.907°; the number of loops is 1 time.

[0174] Furthermore, optimize the final pitch angle to obtain the optimal final pitch angle and construct a loop. Extrapolate without power to the optimal ignition time of 390.653 s for ignition according to the current position speed. Then extrapolate with power. During this period, the pitch angle of the second-stage flight is obtained by interpolation according to the burning time of the second-stage engine after ignition. The optimal yaw angle is 32.907°, and the roll angle of the second-stage flight is 0. Then perform extrapolation with power. When the second-stage engine burns out, continue to extrapolate without power until the Y-axis is less than or equal to the Y-axis of the reentry point (-3062 m) and stop. Calculate that the X-direction position deviation of the reentry point is -1297 m.

[0175] Then design the adjustment step size of the final pitch angle to be 0.1°. Set the final pitch angle to 89.1° after the second-stage engine burns out and regenerate as Figure 3The list of pitch angle attitudes interpolated with the engine combustion time as shown; then extrapolate without power to the ignition time at 390.653 s according to the current position and speed; then extrapolate with power, during which the pitch angle is obtained by interpolation according to the combustion time of the second-stage engine after ignition. After the combustion time enters, the optimal yaw angle is 32.907°, and the rolling angle of the second-stage flight is 0, and extrapolate with power; when the combustion of the second-stage engine is exhausted, continue to extrapolate without power until the Y-axis is less than or equal to the Y-axis of the reentry point (-3062 m) and stop. Calculate that the X-direction position deviation at the reentry point is -1257 m.

[0176] Thus, the calculation obtains The pitch angle attitude after the combustion of the second-stage engine in the next cycle is 82.76°.

[0177] When the number of iterations > 10 times, or the deviation of the X-direction reentry point from the standard orbit constraint < 600 m, jump out of the loop, and obtain the optimal pitch angle attitude after the combustion of the second-stage engine as 82.76°; the number of loops is 1 time.

[0178] When the flight time reaches the optimal ignition time of 390.653 s, perform second-stage ignition, and then fly according to the generated optimal pitch final value, optimal yaw angle, and rolling angle until the reentry point. The actual reentry error is (-543; 2.83; 1862 m); the inclination error at the reentry point is 0.08°.

[0179] In a specific embodiment, assume that the current time of a certain aircraft is 360 s; the current position is (927385; -7599; -42231 m); the current speed is (2796; -1242; -1001 m / s), and the second-stage engine has burned out. It is required to reenter at an angle of attack of α0 = 0 at the reentry point, and calculate the reentry attitude at 0 angle of attack.

[0180] The calculation obtains:

[0181] Dxyzp = 2645491331.02;

[0182] Nxyz = 928377.99

[0183] NVxyz = 3220.15

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

[0185] In a second aspect, the embodiments of the present application further provide an energy planning device for the descending arc section of an aircraft.

[0186] In one embodiment, with reference to Figure 4 , Figure 4This is a schematic diagram of the functional modules of an embodiment of the energy planning device for the descending arc section of the aircraft in this application. As Figure 4 shown, the energy planning device for the descending arc section of the aircraft includes:

[0187] A first calculation module, which is used to calculate the lowest RCS attitude of the aircraft according to 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] A first adjustment module, which is used to adjust the attitude of the aircraft to the lowest RCS attitude and control the aircraft to fly without power when the flight altitude of the aircraft is greater than the preset altitude;

[0189] A second calculation module, which is used to search for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbital inclination constraint and re-entry point coordinate constraint of the re-entry point of the aircraft after the aircraft flies without power to the highest point of the orbit;

[0190] A second adjustment module, which is used to control the second-stage engine of the aircraft to ignite at the optimal ignition time and adjust the attitude of the aircraft to the optimal ignition attitude.

[0191] Further, in an embodiment, the first calculation module is further used for:

[0192] Determine the vector of the aircraft relative to the target detection point according to the current vector position of the aircraft in the launch inertial system and the vector position of the target detection point in the launch inertial system;

[0193] Perform vector normalization on the vector of the aircraft relative to the target detection point to obtain the corresponding first unit vector;

[0194] Perform vector normalization on the current vector position of the aircraft to obtain the corresponding second unit vector;

[0195] Perform a cross product calculation on the first unit vector and the second unit vector in the opposite direction to obtain the reference axis;

[0196] Perform a cross product calculation on the reference axis and the first unit vector to obtain the corrected reference axis;

[0197] Calculate the lowest RCS pitch angle, the lowest RCS yaw angle, and the lowest RCS roll angle in the lowest RCS attitude according to the first unit vector, the corrected reference axis, and the reference axis:

[0198]

[0199] ψ = a sin(-Zx)

[0200] γ = a tan 2(Zy,Zz)

[0201] Where is the lowest RCS pitch angle, ψ 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] Further, in one embodiment, the second calculation module is further configured to:

[0203] Search for the optimal ignition time of the secondary engine of the aircraft according to the orbital inclination constraint of the aircraft;

[0204] Search for the optimal yaw angle of the aircraft after the secondary engine ignites according to the Z - direction coordinate constraint of the re - entry point of the aircraft;

[0205] Search for the final optimal pitch angle of the aircraft after the secondary engine ignites according to the X - direction coordinate constraint of the re - entry point of the aircraft.

[0206] Further, in one embodiment, the second calculation module is further configured to:

[0207] Set the correspondence between the combustion process of the secondary engine and the pitch angle of the aircraft after the secondary engine of the aircraft ignites as:

[0208] During the first one - third of the time of the combustion of the secondary engine of the aircraft, the pitch angle of the aircraft remains the initial pitch angle at the time of ignition of the secondary engine;

[0209] During the middle one - third of the time of the combustion of the secondary engine of the aircraft, the pitch angle of the aircraft is linearly adjusted from the initial pitch angle to the final pitch angle;

[0210] During the last one - third of the time of the combustion of the secondary engine of the aircraft, the pitch angle of the aircraft remains the final pitch angle.

[0211] Further, in one embodiment, the second calculation module is further configured to: cyclically calculate the first orbital inclination corresponding to the aircraft when extrapolated to the re - entry point altitude based on the position and velocity of the aircraft during virtual ignition at the planned ignition time of the secondary engine, the secondary flight pitch angle, secondary flight yaw angle, and secondary flight roll angle of the aircraft after virtual ignition, where the secondary flight pitch angle is obtained by time interpolation according to the correspondence between the combustion process of the secondary engine and the pitch angle, the secondary flight yaw angle is the yaw angle at the virtual ignition time, and the secondary flight roll angle is zero;

[0212] Calculate the inclination error between the first orbital inclination and the preset standard inclination value;

[0213] Add the preset time adjustment step to the planned ignition time to obtain the adjusted ignition time. Based on the position and velocity of the aircraft when the second-stage engine is virtually ignited at the adjusted 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, calculate the second orbital inclination angle when the aircraft is extrapolated to the reentry point altitude.

[0214] Calculate the derivative of the orbital inclination angle with respect to the time adjustment step based on the first orbital inclination angle and the second orbital inclination angle.

[0215] Update the planned ignition time according to the time adjustment step and the ratio of the inclination error to the derivative of the orbital inclination angle with respect to the time adjustment step, and perform iteration with this until the inclination error is less than the preset inclination error threshold or the preset number of iterations is reached, and use the virtual ignition time corresponding to the end of the iteration as the optimal ignition time.

[0216] Further, in one embodiment, the second calculation module is further configured to: repeatedly calculate the first Z coordinate when the aircraft is extrapolated to the reentry point altitude based on the position and velocity of the aircraft 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 of the aircraft after virtual ignition, where the second-stage flight pitch angle of the aircraft after virtual ignition is obtained by time interpolation according to the correspondence between the combustion process of the second-stage engine 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. <>

[0217] Calculate the first Z coordinate error between the first Z coordinate and the preset standard Z coordinate.

[0218] Add the preset yaw angle adjustment step to the second-stage flight yaw angle to obtain the adjusted yaw angle. Based on the position and velocity of the aircraft when the second-stage engine is virtually ignited at the optimal ignition time, the second-stage flight pitch angle, adjusted yaw angle, and second-stage flight roll angle of the aircraft after virtual ignition, calculate the second Z coordinate when the aircraft is extrapolated to the reentry point altitude.

[0219] Calculate the second Z coordinate error between the second Z coordinate and the standard Z coordinate.

[0220] Calculate the derivative of the Z coordinate error with respect to the yaw angle adjustment step based on the first Z coordinate error and the second Z coordinate error.

[0221] Update the second-stage flight yaw angle according to the yaw angle adjustment step and the ratio of the Z coordinate error to the derivative of the Z coordinate error with respect to the yaw angle adjustment step, and perform iteration with this until the Z coordinate error is less than the preset Z coordinate error threshold or the preset number of iterations is reached, and use the second-stage flight yaw angle after virtual ignition corresponding to the end of the iteration as the optimal yaw angle.

[0222] Further, in one embodiment, the second calculation module is further configured to: cyclically calculate the first X - coordinate when the aircraft is extrapolated to the re - entry point altitude based on the position and velocity of the aircraft during virtual ignition at the optimal ignition moment of the secondary engine, the secondary pitch angle, the optimal yaw angle, and the secondary flight roll angle of the aircraft after virtual ignition, wherein the secondary flight pitch angle of the aircraft after virtual ignition is obtained by time interpolation according to the corresponding relationship between the combustion process of the secondary engine and the pitch angle, the final value of the pitch angle in the corresponding relationship is the planned final pitch angle value, and the roll angle is zero;

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

[0224] Add the preset pitch angle final value adjustment step to the planned final pitch angle value to obtain the adjusted final pitch angle value. Calculate the second X - coordinate when the aircraft is extrapolated to the re - entry point altitude based on the position and velocity of the aircraft during virtual ignition at the optimal ignition moment of the secondary engine, the secondary flight pitch angle, the optimal yaw angle, and the secondary flight roll angle of the aircraft after virtual ignition;

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

[0226] Calculate the derivative of the X - direction error with respect to the pitch angle final value adjustment step according to the first X - direction error and the second X - direction error;

[0227] Update the final pitch angle value according to the pitch angle final value adjustment step and the ratio of the X - direction error to the derivative of the X - direction error with respect to the pitch angle final value adjustment step, and perform iteration accordingly until the X - direction error is less than the preset X - direction error threshold or the preset number of iterations is reached, and use the final pitch angle value after virtual ignition corresponding to the end of the iteration as the optimal final pitch angle value.

[0228] Further, in one embodiment, the second calculation module is further configured to: after the secondary engine of the aircraft burns out completely and the altitude of the aircraft is less than the preset altitude, determine the vector of the aircraft relative to the earth's center;

[0229] Calculate the dot product of the vector of the aircraft relative to the earth's center and the velocity vector of the aircraft in the launch inertial system;

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

[0231] Calculate the re - entry pitch angle, re - entry yaw angle, and re - entry roll angle when the re - entry attack angle of the aircraft satisfies the constraint attack angle according to the dot product, the modulus of the vector of the aircraft relative to the earth's center, and the modulus of the velocity vector of the aircraft:

[0232]

[0233] γzr = 0

[0234] Among them, is the reentry pitch angle, ψzr 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 aircraft relative to the earth's center, and NVxyz is the magnitude of the velocity vector of the aircraft.

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

[0236] When receiving the position of a new target detection point, calculate the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the new target detection point.

[0237] Among them, the function implementation of each module in the above aircraft descending arc section energy planning device corresponds to each step in the above aircraft descending arc section energy planning method embodiment, and its function and implementation process will not be elaborated here one by one.

[0238] It should be noted that the serial numbers of the above embodiments of the present application are only for description and do not represent the advantages and disadvantages of the embodiments.

[0239] The terms "including" and "having" and any variations thereof in the specification and claims of the present application and the above drawings are intended to cover non-exclusive inclusion. For example, a process, method, system, product or device that includes a series of steps or units is not limited to the listed steps or units, but optionally further includes steps or units not listed, or optionally further includes other steps or units inherent to these processes, methods, products or devices. The descriptions of the terms "first", "second", "third", etc. are used to distinguish different objects, etc., and do not represent a sequence, nor do they limit that "first", "second", and "third" are different types.

[0240] In the description of the embodiments of the present application, "exemplary", "for example" or "for instance" etc. are used to represent examples, illustrations or explanations. Any embodiment or design solution described as "exemplary", "for example" or "for instance" in the embodiments of the present application should not be construed as being more preferred or having more advantages than other embodiments or design solutions. Rather, the use of the words "exemplary", "for example" or "for instance" etc. is intended to present the relevant concepts in a specific manner.

[0241] In the description of the embodiments of the present application, unless otherwise specified, " / " means "or". For example, A / B can represent A or B; "and / or" in the text is only a description of the association relationship of the associated objects, indicating that there can be three relationships. For example, A and / or B can represent: A exists alone, A and B exist simultaneously, and B exists alone. In addition, in the description of the embodiments of the present application, "a plurality of" means two or more than two.

[0242] In some of the processes described in the embodiments of the present application, a plurality of operations or steps appear in a specific order. However, it should be understood that these operations or steps may not be executed in the order in which they appear in the embodiments of the present application or may be executed in parallel. The serial numbers of the operations are only used to distinguish the different operations, and the serial numbers themselves do not represent any order of execution. In addition, these processes may include more or fewer operations, and these operations or steps may be executed in sequence or in parallel, and these operations or steps may be combined.

[0243] Through the description of the above embodiments, those skilled in the art can clearly understand that the above-described embodiment methods can be implemented by means of software plus a necessary general hardware platform. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation. Based on such an understanding, the technical solution of the present 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 for causing a terminal device to execute the methods described in the various embodiments of the present application.

[0244] The above are only the preferred embodiments of the present application, and do not limit the patent scope of the present application. Any equivalent structure or equivalent process transformation made by using the content of the specification and drawings of the present application, or directly or indirectly applied in other related technical fields, shall be equally included in the patent protection scope of the present application.

Claims

1. An energy planning method for the descending arc section of an aircraft, characterized in that, The energy planning method for the descending arc section of the aircraft includes: After the separation of the first-stage engine of the aircraft, calculate the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the target detection point; When the flight altitude of the aircraft is greater than the preset altitude, adjust the attitude of the aircraft to the lowest RCS attitude and control the aircraft to fly without power; After the aircraft flies without power to the highest point of the orbit, search for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbit inclination constraint and re-entry point coordinate constraint of the re-entry point of the aircraft; Control the second-stage engine of the aircraft to ignite at the optimal ignition time and adjust the attitude of the aircraft to the optimal ignition attitude.

2. The energy planning method for the descending arc section of an aircraft as described in claim 1, characterized in that, The calculation of the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the target detection point includes: Determine the vector of the aircraft relative to the target detection point according to the current vector position of the aircraft in the launch inertial system and the vector position of the target detection point in the launch inertial system; Perform vector normalization on the vector of the aircraft relative to the target detection point to obtain the corresponding first unit vector; Perform vector normalization on the current vector position of the aircraft to obtain the corresponding second unit vector; Perform a cross product calculation on the first unit vector and the second unit vector in the opposite direction to obtain the reference axis; Perform a cross product calculation on the reference axis and the first unit vector to obtain the corrected reference axis; Calculate the lowest RCS pitch angle, lowest RCS yaw angle, and lowest RCS roll angle in the lowest RCS attitude according to the first unit vector, the corrected reference axis, and the reference axis: ψ = a sin(-Zx) γ = a tan2(Zy, Zz) Among them, is the lowest RCS pitch angle, ψ 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 descending arc section of an aircraft as described in claim 1, wherein, The search for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbit inclination constraint and re-entry point coordinate constraint of the re-entry point of the aircraft includes: Search for the optimal ignition time of the second-stage engine of the aircraft according to the orbit inclination constraint of the aircraft; Search for the optimal yaw angle of the aircraft after the second-stage engine ignition according to the Z-direction coordinate constraint of the re-entry point of the aircraft; Search for the final optimal pitch angle of the aircraft after the second-stage engine ignition according to the X-direction coordinate constraint of the re-entry point of the aircraft.

4. The method for planning energy during the descent arc of an aircraft according to claim 3, wherein: Before searching for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbit inclination constraint and re-entry point coordinate constraint of the re-entry point of the aircraft, it also includes: Set the corresponding relationship between the combustion process of the second-stage engine and the pitch angle of the aircraft after the second-stage engine ignition of the aircraft as: During the first one-third of the time of the combustion of the second-stage engine of the aircraft, the pitch angle of the aircraft remains the initial pitch angle at the time of the second-stage engine ignition; During the middle one-third of the time of the combustion of the second-stage engine of the aircraft, the pitch angle of the aircraft is linearly adjusted from the initial pitch angle to the final pitch angle; During the last one-third of the time of the combustion of the second-stage engine of the aircraft, the pitch angle of the aircraft remains the final pitch angle.

5. The energy planning method for the descending arc section of an aircraft according to claim 4, characterized in that The search for the optimal ignition time of the second-stage engine of the aircraft according to the orbit inclination constraint of the aircraft includes: Loop based on the position and velocity of the aircraft at the planned ignition time of the second-stage engine for virtual ignition, the pitch angle, yaw angle, and roll angle of the second-stage flight of the aircraft after virtual ignition, and calculate the first orbital inclination angle corresponding to the aircraft extrapolated to the altitude of the reentry point. Among them, the pitch angle of the second-stage flight is obtained by time interpolation according to the corresponding relationship between the combustion process of the second-stage engine and the pitch angle. The yaw angle of the second-stage flight is the yaw angle at the virtual ignition time, and the roll angle of the second-stage flight is zero; Calculate the inclination error between the first orbital inclination angle and the preset standard inclination value; Add the preset time adjustment step size to the planned ignition time to obtain the adjusted ignition time. Based on the position and velocity of the aircraft at the adjusted ignition time of the second-stage engine for virtual ignition, the pitch angle, yaw angle, and roll angle of the second-stage flight of the aircraft after virtual ignition, calculate the second orbital inclination angle when the aircraft is extrapolated to the altitude of the reentry point; Calculate the derivative of the orbital inclination angle with respect to the time adjustment step size based on the first orbital inclination angle and the second orbital inclination angle; Update the planned ignition time according to 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, and perform iteration with this until the inclination error is less than the preset inclination error threshold or the preset iteration number is reached, and use the virtual ignition time corresponding to the end of the iteration as the optimal ignition time.

6. The energy planning method for the descending arc section of an aircraft according to claim 4, characterized in that The searching for the optimal yaw angle of the aircraft after the second-stage engine ignition according to the Z-direction coordinate constraint of the reentry point of the aircraft includes: Loop based on the position and velocity of the aircraft at the optimal ignition time of the second-stage engine for virtual ignition, the pitch angle, yaw angle, and roll angle of the second-stage flight of the aircraft after virtual ignition, and calculate the first Z-direction coordinate when the aircraft is extrapolated to the altitude of the reentry point. Among them, the pitch angle of the second-stage flight of the aircraft after virtual ignition is obtained by time interpolation according to the corresponding relationship between the combustion process of the second-stage engine and the pitch angle. The yaw angle of the second-stage flight in the first iteration is the yaw angle at the optimal ignition time, and the roll angle of the second-stage flight is zero; Calculate the first Z-direction error between the first Z-direction coordinate and the preset standard Z-direction coordinate; Add the preset yaw angle adjustment step size to the yaw angle of the second-stage flight to obtain the adjusted yaw angle. Based on the position and velocity of the aircraft at the optimal ignition time of the second-stage engine for virtual ignition, the pitch angle, adjusted yaw angle, and roll angle of the second-stage flight of the aircraft after virtual ignition, calculate the second Z-direction coordinate when the aircraft is extrapolated to the altitude of the reentry point; Calculate the second Z-direction error between the second Z-direction coordinate and the standard Z-direction coordinate; Calculate the derivative of the Z-direction error with respect to the yaw angle adjustment step size based on the first Z-direction error and the second Z-direction error; Update the yaw angle of the second-stage flight according to the yaw angle adjustment step size and the ratio of the Z-direction error to the derivative of the Z-direction error with respect to the yaw angle adjustment step size, and perform iteration with this until the Z-direction error is less than the preset Z-direction error threshold or the preset iteration number is reached, and use the yaw angle of the second-stage flight after virtual ignition corresponding to the end of the iteration as the optimal yaw angle.

7. The energy planning method for the descending arc section of an aircraft according to claim 4, wherein Searching for the optimal final pitch angle of the aircraft after the second-stage engine ignition according to the reentry point X-direction coordinate constraint of the aircraft, including: Circulating according to the position and velocity of the aircraft during virtual ignition at the optimal ignition moment of the second-stage engine, the corresponding second-stage pitch angle of the aircraft after virtual ignition, the optimal yaw angle, and the second-stage flight roll angle, calculating the first X-direction coordinate when the aircraft is extrapolated to the reentry point altitude, where the second-stage flight pitch angle of the aircraft after virtual ignition is obtained by time interpolation according to the corresponding relationship between the second-stage engine combustion process and the pitch angle. In the corresponding relationship, the final pitch angle value is the planned final pitch angle value, and the roll angle is zero; Calculating the first X-direction error between the first X-direction coordinate and the preset standard X-direction coordinate; Adding the preset final pitch angle adjustment step to the planned final pitch angle value to obtain the adjusted final pitch angle value. According to the position and velocity of the aircraft during virtual ignition at the optimal ignition moment of the second-stage engine, the second-stage flight pitch angle, the optimal yaw angle, and the second-stage flight roll angle of the aircraft after virtual ignition, calculating the second X-direction coordinate when the aircraft is extrapolated to the reentry point altitude; Calculating the second X-direction error between the second X-direction coordinate and the standard X-direction coordinate; Calculating the derivative of the X-direction error with respect to the final pitch angle adjustment step according to the first X-direction error and the second X-direction error; Updating the final pitch angle value according to the final pitch angle adjustment step and the ratio of the X-direction error to the derivative of the X-direction error with respect to the final pitch angle adjustment step, and performing iteration accordingly until the X-direction error is less than the preset X-direction error threshold or the preset iteration times are reached, and taking the final pitch angle value after virtual ignition corresponding to the end of the iteration as the optimal final pitch angle value.

8. The energy planning method for the descending arc section of an aircraft as described in claim 1, characterized in that, This method further includes: After the second-stage engine of the aircraft burns out completely and the altitude of the aircraft is less than the preset altitude, determining the vector of the aircraft relative to the earth's center; Calculating the dot product of the vector of the aircraft relative to the earth's center and the velocity vector of the aircraft in the launch inertial system; Determining the modulus of the vector of the aircraft relative to the earth's center and the modulus of the velocity vector of the aircraft; Calculating the reentry pitch angle, reentry yaw angle, and reentry roll angle when the reentry attack angle of the aircraft satisfies the constraint attack angle according to the dot product, the modulus of the vector of the aircraft relative to the earth's center, and the modulus of the velocity vector of the aircraft: γzr = 0 wherein, is the reentry pitch angle, ψzr 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 aircraft relative to the center of the earth, and NVxyz is the magnitude of the velocity vector of the aircraft.

9. The energy planning method for the descending arc section of an aircraft according to claim 1, characterized in that After the first-stage engine of the aircraft separates, calculating the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the target detection point, including: When receiving the position of the new target detection point, calculating the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the new target detection point.

10. An energy planning device for the descending arc section of an aircraft, characterized in that, The aircraft descending arc section energy planning device includes: A first calculation module, which is used to calculate the lowest RCS attitude of the aircraft according to the current position of the aircraft and the position of the target detection point after the first-stage engine of the aircraft separates; A first adjustment module, which is used to adjust the attitude of the aircraft to the lowest RCS attitude and control the aircraft to fly without power after the flight altitude of the aircraft is greater than the preset altitude; A second calculation module, which is used to search for the optimal ignition time and optimal ignition attitude of the second-stage engine of the aircraft according to the orbital inclination constraint and the re-entry point coordinate constraint of the re-entry point of the aircraft after the aircraft flies without power to the highest point of the orbit; A second adjustment module, which is used to control the second-stage engine of the aircraft to ignite at the optimal ignition time and adjust the attitude of the aircraft to the optimal ignition attitude.

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