Guidance method for compound guided vehicle under satellite denial condition

Abstract

The application discloses a kind of satellite rejection condition under the guidance method of composite guided vehicle, in this method, vehicle is in middle guidance section, based on satellite signal and attitude sensing system obtains the expected acceleration of vehicle, according to this control vehicle flies to target, in this process, if encounter satellite rejection, based on the expected acceleration of vehicle obtained by satellite signal applied at last time, until real-time satellite signal is reacquired, after the launch scheduled time of vehicle, laser seeker is started;When laser seeker captures target, the expected acceleration of vehicle is obtained in real time through new line-of-sight angle constraint guidance law, generates rudder instruction based on the expected acceleration of vehicle, controls rudder machine to work, controls vehicle to fly to target, and collides target with expected line-of-sight angle, through the new line-of-sight angle constraint guidance law compensation correction deviation caused by satellite rejection, finally makes vehicle hit target.

Description

Technical Field

[0001] This invention relates to aircraft guidance and control methods, specifically to a guidance method for a composite guidance aircraft under satellite denial conditions. Background Technology

[0002] A composite-guided aircraft refers to an aircraft whose flight trajectory employs different guidance methods in the same or different phases (such as midcourse or terminal phase). A satellite-plus-laser composite-guided aircraft utilizes a satellite navigation system for midcourse guidance, controlling the aircraft to complete autonomous flight and approach the target area during the midcourse phase; in the terminal guidance phase, laser guidance is used to track and strike the target. Compared to traditional single-mode guided aircraft, satellite-plus-laser composite-guided aircraft significantly improve the accuracy of mid-to-long-range guided aircraft.

[0003] However, due to the fragility of satellite signals, when an aircraft is in a harsh environment, satellite signals can be blocked or altered at any time, resulting in satellite denial. Under satellite denial conditions, composite guidance aircraft using satellite navigation systems in the mid-course guidance phase will experience significant deviations. To ensure the aircraft accurately hits the target, it is necessary to correct the errors generated in the mid-course guidance phase in the terminal guidance phase. However, current technologies lack effective correction schemes for aircraft entering the terminal guidance phase with large errors. Therefore, there is an urgent need for a method that can correct the large deviations caused by satellite denial and other factors in the mid-course guidance phase during the terminal guidance phase, ultimately enabling the aircraft to maintain high strike accuracy. This is of great significance for improving the aircraft's anti-jamming capabilities.

[0004] Based on the above problems, the inventors have conducted in-depth research on a composite guidance method using laser guidance in the terminal guidance phase, in order to design a spacecraft control method with strong correction capabilities that can still ensure hit accuracy in the event of satellite denial. Summary of the Invention

[0005] To overcome the aforementioned problems, the inventors conducted in-depth research and designed a guidance method for a composite-guided aircraft under satellite denial conditions. In this method, during the mid-guidance phase, the aircraft obtains its desired acceleration based on satellite signals and an attitude sensing system, and controls the aircraft to fly towards the target accordingly. If satellite denial is encountered during this process, the aircraft obtains its desired acceleration based on the satellite signals applied at the previous moment until real-time satellite signals are regained. After a predetermined launch time, the laser seeker is activated. Once the laser seeker acquires the target, it obtains the aircraft's desired acceleration in real time through a novel line-of-sight constraint guidance law. Based on the aircraft's desired acceleration, rudder commands are generated to control the rudder mechanism to operate, guiding the aircraft to fly towards the target and colliding with the target at the desired line-of-sight angle. The novel line-of-sight constraint guidance law compensates for and corrects the deviation caused by satellite denial, ultimately enabling the aircraft to hit the target, thus completing the invention.

[0006] Specifically, the purpose of this invention is to provide a guidance method for a composite guidance vehicle under satellite denial conditions, the method comprising the following steps:

[0007] Step 1: During the mid-course guidance phase, the aircraft obtains the desired acceleration based on satellite signals and the attitude sensing system, and controls the aircraft to fly toward the target accordingly. After the predetermined launch time, the laser seeker is activated.

[0008] Step 2: After the laser seeker acquires the target, the desired acceleration of the aircraft is obtained in real time through the new line-of-sight angle constraint guidance law. Based on the desired acceleration of the aircraft, rudder commands are generated to control the rudder motor to operate, control the aircraft to fly towards the target, and collide with the target at the desired line-of-sight angle.

[0009] In step 1, when the satellite denies the signal, the desired acceleration of the spacecraft is obtained based on the satellite signal applied at the previous moment until the real-time satellite signal is obtained again.

[0010] In step 2, the novel line-of-sight angle constraint guidance law obtains the desired acceleration of the aircraft in real time using the following equation (i):

[0011]

[0012] Among them, a M This represents the desired acceleration of the aircraft.

[0013] r represents the distance between the aircraft and the target, which is calculated using ephemeris data of the aircraft and the target;

[0014] This represents the component of the relative velocity between the aircraft and the target along the line of sight.

[0015] Indicates the angular velocity of the line of sight of the projectile.

[0016] θ represents the projectile-to-eye line of sight angle, obtained through the laser seeker.

[0017] θ d Indicates the desired viewing angle.

[0018] This represents the angle between the aircraft's velocity vector and the line of sight.

[0019] M represents a constant greater than 0.

[0020] S represents the sliding surface.

[0021] β represents a constant greater than 0.

[0022] The beneficial effects of this invention include:

[0023] (1) The guidance method of the composite guidance vehicle under satellite denial conditions provided by the present invention has extremely high universality. It can correct the flight trajectory when there is a deviation in the mid-guidance phase, so that the vehicle can finally hit the target. As a result, it eliminates the adverse effects of external factors such as satellite signal interference on the hit accuracy.

[0024] (2) The guidance method of the composite guided aircraft under satellite denial conditions provided by the present invention, wherein the desired acceleration of the aircraft is obtained in real time by a novel line-of-sight angle constraint guidance law, thereby enabling the aircraft to hit the target at the desired line-of-sight angle. At the desired line-of-sight angle, the aircraft can hit the weak point of the target at a specific angle, thereby obtaining the best hit effect.

[0025] (3) The guidance method for a composite guided aircraft under satellite denial conditions provided by the present invention can specifically solve the problem of large deviation of the aircraft at the beginning of terminal guidance, and will not cause the aircraft to become unstable or inaccurate due to the short-term required overload exceeding the actual working range of the servo motor; that is, the new line-of-sight angle constraint guidance law provides a continuous and smooth required overload through the sliding mode surface; correspondingly, the traditional proportional guidance guidance law is a guidance method that makes the rotational angular velocity of the aircraft velocity vector proportional to the rotational angular velocity of the line of sight. If there is a large deviation in the initial stage of terminal guidance, in order to achieve the constraint effect, the required normal overload will inevitably be less than the available normal overload, which will eventually lead to the aircraft becoming unstable or inaccurate. Attached Figure Description

[0026] Figure 1 This paper presents a schematic diagram of the overall logic of the guidance method for a composite guidance vehicle under satellite denial conditions.

[0027] Figure 2 A schematic diagram of the flight trajectory of the aircraft in the embodiment is shown;

[0028] Figure 3 A schematic diagram showing the change of the aircraft's line-of-sight angle over time in the embodiment is shown. Detailed Implementation

[0029] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present invention will become clearer and more apparent.

[0030] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0031] According to the present invention, a guidance method for a composite guidance vehicle under satellite denial conditions is provided, such as... Figure 1 As shown, the method includes the following steps:

[0032] Step 1: During the mid-course guidance phase, the aircraft obtains the desired acceleration based on satellite signals and the attitude sensing system, and controls the aircraft to fly toward the target accordingly. After the predetermined launch time, the laser seeker is activated. Before the launch, the flight time of the aircraft is estimated based on the target position information, the launch position of the aircraft, and the flight speed of the aircraft. The aircraft is set to activate the laser seeker to enter the terminal guidance phase when it is about 4 kilometers away from the target.

[0033] In this application, before the launch of the aircraft, the ephemeris information of the aircraft and the target, as well as the preset expected flight time and expected line-of-sight information are installed in the aircraft. After launch, the aircraft begins the program setup before the highest point and enters the mid-course guidance phase.

[0034] During the mid-course guidance phase, when a satellite denies the mission, the spacecraft obtains its desired acceleration based on the satellite signal applied in the previous moment, until it regains real-time satellite signal.

[0035] In this application, satellite denial means that the aircraft cannot receive signals from the satellite and cannot update its own status information based on real-time satellite signals;

[0036] Preferably, during the mid-course guidance phase, the spacecraft remains in a satellite denial state for 16 consecutive seconds, resulting in a relatively small directional error. This allows the laser seeker to promptly capture the diffusely reflected laser signal from the target after activation, and correct it using a novel line-of-sight angle constraint guidance law in the subsequent terminal guidance phase, ultimately enabling the spacecraft to hit the target.

[0037] In this application, the attitude sensing system uses a geomagnetic sensor to measure the roll angle and rotational speed of the aircraft. Compared with traditional space-oriented gyroscopes, the geomagnetic sensor is not affected by the frame angle constraint, so it can work normally under high overload conditions.

[0038] Step 2: After the laser seeker acquires the target, the desired acceleration of the aircraft is obtained in real time through the new line-of-sight angle constraint guidance law. Based on the desired acceleration of the aircraft, rudder commands are generated to control the rudder motor to operate, control the aircraft to fly towards the target, and collide with the target at the desired line-of-sight angle.

[0039] In this application, the aircraft is equipped with a servo system, which generates rudder deflection commands based on the aircraft's desired acceleration, thereby controlling the servo deflection. It can also feed back the rudder deflection status to the microprocessor that calculates the desired acceleration, perform feedforward compensation on the rudder deflection commands, and improve rudder steering accuracy.

[0040] In this application, a strapdown laser seeker is selected as the laser seeker, which is responsible for measuring the line-of-sight angle. Compared with traditional platform seekers, strapdown seekers do not need to be installed on the platform and can be directly fixed to the aircraft, which not only saves the airframe loading space, but also performs well under high overload conditions.

[0041] Once the laser seeker captures the laser echo reflected from the target, it can be considered that the target has been captured, and at this time it enters the terminal guidance phase.

[0042] In step 2, the novel line-of-sight angle constrained guidance law obtains the desired acceleration of the aircraft in real time using the following equation (i):

[0043]

[0044] Among them, a M This represents the desired acceleration of the aircraft.

[0045] r represents the distance between the aircraft and the target, which is calculated using ephemeris data of the aircraft and the target;

[0046] This represents the line-of-sight component of the relative velocity between the aircraft and the target, which can be obtained via satellite signals;

[0047] In this application, under normal circumstances, the aforementioned r and [other parameters] are still obtained via satellite signals during the terminal guidance phase. When encountering satellite denial, r and [other information] are obtained through the laser seeker. In this application, the laser signal illuminating the target is an intermittent signal, which can be emitted by a third party, i.e., illuminating once every predetermined time interval t1, and the duration of each illumination is also constant. The aircraft pre-stores the duration and interval t1 of the illumination. The time elapsed from the i-th laser emission to the receipt of the diffusely reflected laser is t. ki The speed of light is c, the distance between the laser illuminator and the target is s, and the speed of the aircraft is V. M r i Let represent the distance between the aircraft and the target measured in the i-th time.

[0048] r i =ct ki -s

[0049] r i+1 =ct k(i+1) -s

[0050] Because the time interval is small, the component of the relative velocity between the aircraft and the target measured in the i-th time along the line of sight is considered as:

[0051]

[0052] Preferably, an active laser illuminator can also be installed on the laser seeker head of the aircraft to actively illuminate the target, thereby obtaining the corresponding r and

[0053] The line-of-sight angular velocity of the projectile is obtained by the following formula:

[0054] θ represents the projectile-to-eye line of sight angle, obtained through the laser seeker.

[0055] θ d Indicates the desired viewing angle.

[0056] This represents the angle between the aircraft's velocity vector and the line of sight, and is simply referred to as the aircraft's velocity vector lead angle.

[0057] M represents a constant greater than 0, preferably 3000.

[0058] S represents the sliding surface.

[0059] β represents a constant greater than 0, and preferably has a value of 1.5.

[0060] In a preferred embodiment, the sliding surface S is obtained by the following formula (ii):

[0061]

[0062] Where t represents the time elapsed after the start of terminal guidance, with the timing starting when the laser seeker captures the diffusely reflected laser signal from the target.

[0063] Preferably, By applying the formula obtained from equation (iii) Differentiation yields:

[0064]

[0065] Where, a′ M This represents the desired acceleration of the aircraft at the previous moment. When it just enters the terminal guidance phase and first uses the new line-of-sight angle constraint guidance law to obtain the desired acceleration of the aircraft, this a′... M The value is the expected acceleration at the last moment of the mid-course guidance phase. In the aircraft of this application, the aircraft is guided and controlled by a conventional proportional guidance law during the mid-course guidance phase.

[0066] V M Indicates the speed of the aircraft.

[0067] Indicates the previous moment When the aircraft first entered the terminal guidance phase and used the new line-of-sight angle constraint guidance law to obtain the desired acceleration, The value of is 0.

[0068] Example

[0069] Three identical aircraft were set up, each with an initial velocity of 800 m / s. The target was 30 km from the launch position, and the environmental disturbances experienced by the three aircraft were exactly the same. Each aircraft ejected a laser fairing 110 seconds after launch and acquired the target via a laser seeker. This was achieved by providing simulated laser signals to the three aircraft via a simulation system 110 seconds after launch.

[0070] The first spacecraft operates under normal conditions, meaning it receives satellite signals in real time during both the mid-course and terminal guidance phases, and there is no satellite rejection. The control process for this spacecraft includes the following steps:

[0071] Step 1: During the mid-course guidance phase, the aircraft obtains its desired acceleration based on satellite signals and the attitude sensing system. Based on this, the aircraft is controlled to fly towards the target. 110 minutes after launch, the laser seeker is activated to receive simulated laser signals. Step 2: Once the laser seeker acquires the target (i.e., after receiving the simulated laser signal), it obtains the aircraft's desired acceleration in real time using a novel line-of-sight constraint guidance law. Based on the aircraft's desired acceleration, rudder commands are generated to control the rudder mechanism, guiding the aircraft towards the target and colliding with it at the desired line-of-sight angle.

[0072] The novel line-of-sight angle constrained guidance law obtains the desired acceleration of the aircraft in real time through the following equation (I):

[0073]

[0074] Among them, a M This represents the desired acceleration of the aircraft.

[0075] r represents the distance between the aircraft and the target;

[0076] This represents the component of the relative velocity between the aircraft and the target along the line of sight.

[0077] Indicates the angular velocity of the line of sight of the projectile.

[0078] θ represents the line-of-sight angle of the bullet.

[0079] θ d This represents the desired viewing angle, with a value of 70°.

[0080] This represents the angle between the aircraft's velocity vector and the line of sight.

[0081] M represents a constant with a value of 3000.

[0082] S represents the sliding surface.

[0083] β represents a constant with a value of 1.5.

[0084] The sliding surface S is obtained by the following formula (ii):

[0085]

[0086] Where t represents the time starting 110 seconds after the spacecraft launches;

[0087] By applying the formula obtained from equation (iii) Differentiation yields:

[0088]

[0089] Where, a′ M This represents the expected acceleration of the aircraft at the previous moment;

[0090] V M Indicates the speed of the aircraft.

[0091] Indicates the previous moment The initial value is 0.

[0092] Ultimately, the first aircraft stably hit the target at the desired line-of-sight angle.

[0093] The second spacecraft falls under the correction scenario, meaning that satellite denial occurs during both the mid-course and terminal guidance phases, specifically between seconds 90-100 and 105-110. During these periods, no simulated satellite signals are provided to the spacecraft. The control process for this second spacecraft includes the following steps:

[0094] Step 1: During the mid-course guidance phase, the aircraft obtains its desired acceleration based on satellite signals and the attitude sensing system, and controls the aircraft to fly toward the target accordingly. 110 minutes after launch, the laser seeker is activated to receive simulated laser signals. During this process, if the satellite denies the target, the aircraft obtains its desired acceleration based on the satellite signals applied at the previous moment, until real-time satellite signals are obtained again.

[0095] Step 2: After the laser seeker acquires the target and receives the simulated laser signal, it obtains the aircraft's desired acceleration in real time through a new line-of-sight angle constraint guidance law. Based on the aircraft's desired acceleration, it generates rudder commands to control the rudder motors to operate, control the aircraft to fly towards the target, and collide with the target at the desired line-of-sight angle.

[0096] The novel line-of-sight angle constrained guidance law obtains the desired acceleration of the aircraft in real time through the following equation (I):

[0097]

[0098] Among them, a M This represents the desired acceleration of the aircraft.

[0099] r represents the distance between the aircraft and the target;

[0100] This represents the component of the relative velocity between the aircraft and the target along the line of sight.

[0101] Indicates the angular velocity of the line of sight of the projectile.

[0102] θ represents the line-of-sight angle of the bullet.

[0103] θ d This represents the desired viewing angle, with a value of 70°.

[0104] This represents the angle between the aircraft's velocity vector and the line of sight.

[0105] M represents a constant with a value of 3000.

[0106] S represents the sliding surface.

[0107] β represents a constant with a value of 1.5.

[0108] When the satellite denies access

[0109] r i =ct ki -s

[0110]

[0111] r i+1 =ct k(i+1) -s

[0112] r i This represents the distance between the aircraft and the target measured in the i-th time.

[0113] Let represent the component of the relative velocity between the aircraft and the target measured in the i-th time along the line of sight:

[0114] r i+1 This represents the distance between the aircraft and the target measured in the (i+1)th time.

[0115] c represents the speed of light.

[0116] t kiThis represents the time elapsed from the i-th laser emission to the receipt of the diffusely reflected laser light.

[0117] s represents the distance between the laser irradiator and the target.

[0118] The sliding surface S is obtained by the following formula (ii):

[0119]

[0120] Where t represents the time starting 110 seconds after the spacecraft launches;

[0121] By applying the formula obtained from equation (iii) Differentiation yields:

[0122]

[0123] Where, a′ M This represents the expected acceleration of the aircraft at the previous moment;

[0124] V M Indicates the speed of the aircraft.

[0125] Indicates the previous moment The initial value is 0.

[0126] Ultimately, the second aircraft was still able to hit the target at the desired line-of-sight angle.

[0127] The third spacecraft is in an abnormal situation, specifically experiencing satellite rejection during its mid-course guidance phase, between seconds 90 and 100. During this period, no simulated satellite signals are provided to the spacecraft. The control process for this third spacecraft includes the following steps:

[0128] Step 1: During the mid-course guidance phase, the aircraft obtains its desired acceleration based on satellite signals and the attitude sensing system, and controls the aircraft to fly toward the target accordingly. 110 minutes after launch, the laser seeker is activated to receive simulated laser signals. During this process, if the satellite denies the target, the aircraft obtains its desired acceleration based on the satellite signals applied at the previous moment, until real-time satellite signals are obtained again.

[0129] Step 2: After the laser seeker acquires the target, i.e., after receiving the simulated laser signal, it uses the proportional guidance rate... To obtain the desired acceleration of the aircraft in real time; N is 4, V and All are obtained through satellite signals;

[0130] Ultimately, the third aircraft failed to hit its target.

[0131] The flight trajectories of the three aircraft mentioned above are as follows: Figure 2 As shown in the figure, the curves of the line-of-sight angle of the three aircraft changing over time are as follows: Figure 3 As shown, according to the above Figure 2 The results shown indicate that under normal circumstances, the guidance method of the composite guidance vehicle under satellite denial conditions enables the vehicle to consistently hit the target. However, under abnormal circumstances, i.e., satellite denial, the vehicle deviates during the mid-course guidance phase, and the proportional guidance rate control method ultimately fails to accurately hit the target. In the correction phase, the guidance method of the composite guidance vehicle under satellite denial conditions is used to correct the deviations caused by satellite denial, ultimately resulting in an accurate target hit. Based on... Figure 3 It can be seen that, regardless of whether there is satellite denial, the guidance method of the composite guidance vehicle under the satellite denial condition can be used to guide and control the vehicle so that the vehicle can hit the target at the desired line-of-sight angle.

[0132] The present invention has been described above with reference to preferred embodiments; however, these embodiments are merely exemplary and illustrative. Various substitutions and modifications can be made to the present invention based on these embodiments, all of which fall within the scope of protection of the present invention.

Claims

1. A guidance method for a composite-guided aircraft under satellite denial conditions, characterized in that, The method includes the following steps: Step 1: During the mid-course guidance phase, the aircraft obtains the desired acceleration based on satellite signals and the attitude sensing system, and controls the aircraft to fly toward the target accordingly. After the predetermined launch time, the laser seeker is activated. Step 2: After the laser seeker acquires the target, it obtains the aircraft's desired acceleration in real time through the line-of-sight angle constraint guidance law. Based on the aircraft's desired acceleration, it generates rudder commands to control the servo motors to operate, guiding the aircraft towards the target and colliding with it at the desired line-of-sight angle. In step 2, the line-of-sight angle constraint guidance law obtains the desired acceleration of the aircraft in real time using the following equation (i): (one) in, This represents the desired acceleration of the aircraft. Indicates the distance between the aircraft and the target. This represents the component of the relative velocity between the aircraft and the target along the line of sight. Indicates the angular velocity of the line of sight of the projectile. Indicates the line-of-sight angle of the bullet. Indicates the desired viewing angle. This represents the angle between the aircraft's velocity vector and the line of sight. Represents a constant greater than 0. Indicates the sliding surface. Represents a constant greater than 0; The sliding surface We obtain it through the following formula (II): (two) in, This indicates the time elapsed since the start of terminal guidance; In step 2, when the satellite denies the request, the distance between the spacecraft and the target is obtained using the following equation (iv). : (Four) in, Indicates the first The distance between the aircraft and the target was measured this time. Represents the speed of light. Indicates from the first The time elapsed from the emission of a laser to the receipt of that diffusely reflected laser. Indicates the distance between the laser illuminator and the target; In step 2, when the satellite denies the target, the line-of-sight component of the relative velocity between the aircraft and the target is obtained using the following equation (v). : (five) in, Indicates the first The distance between the aircraft and the target was measured this time. Indicates the time interval between two laser irradiations. ， From the The time elapsed from the emission of a laser to the receipt of that diffusely reflected laser is , Indicates the first The distance between the aircraft and the target was measured.

2. The guidance method for a composite guidance vehicle under satellite denial conditions according to claim 1, characterized in that, In step 1, when the satellite denies the signal, the desired acceleration of the spacecraft is obtained based on the satellite signal applied at the previous moment until the real-time satellite signal is regained.

3. The guidance method for a composite guidance vehicle under satellite denial conditions according to claim 1, characterized in that, By applying the formula obtained from equation (iii) Differentiation yields: (three) in, This represents the expected acceleration of the aircraft at the previous moment. Indicates the speed of the aircraft. Indicates the previous moment .

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