An aircraft photoelectric guidance method based on double-channel coupling
By using a dual-channel coupled photoelectric guidance method to process and compensate for the aircraft's pitch and azimuth signals, the problem of low landing accuracy caused by pitch and yaw channel coupling was solved, and high-precision and stable landing of the aircraft was achieved.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- CHINESE PEOPLES LIBERATION ARMY NAVAL SPECIALTY MEDICAL CENT
- Filing Date
- 2022-12-09
- Publication Date
- 2026-07-21
AI Technical Summary
Traditional aircraft carrier landing control methods suffer from poor landing accuracy due to severe coupling between pitch and yaw channels.
A dual-channel coupled electro-optical guidance method is adopted. The pitch and azimuth signals of the aircraft are measured by the electro-optical guidance system. Combined with inertial differential hybrid and adaptive law design, the pitch and azimuth signals are integrated and compensated to eliminate the disturbances caused by channel coupling.
It improves the accuracy and stability of aircraft landing, and is particularly suitable for aircraft control with strong nonlinearity and severe dual-channel coupling, and has high engineering practical value.
Smart Images
Figure CN115903863B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aircraft recovery and position control, and more specifically, to an aircraft landing control and guidance method employing photoelectric guidance. Background Technology
[0002] With the development of optoelectronic technology and the wide application of aircraft in both military and civilian fields, guidance methods related to optoelectronic-assisted aircraft landing and recovery have attracted widespread interest and attention from engineers and aircraft research enthusiasts worldwide. However, due to the complexity of precise aerodynamic modeling for low-altitude flight and the inherent severe coupling between pitch and yaw channels, traditional carrier landing control is based on linear decoupling. This means the pitch channel does not require any information from the yaw channel, and vice versa; the two channels are completely isolated and designed independently. While this method is simple, it essentially severs the coupling between the two channels, leading to significant challenges in heavily coupled aircraft. Based on this background, this invention proposes a dual-channel coupling method, designing both channels holistically. Furthermore, the information from the two channels is fused and fed back to form the control law, achieving good results in experiments and demonstrating the high engineering practical value of this invention.
[0003] It should be noted that the information in the background section above is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those skilled in the art. Summary of the Invention
[0004] The purpose of this invention is to provide an electro-optical guidance landing method for aircraft based on dual-channel coupling, thereby overcoming the problem of low landing accuracy caused by severe coupling of pitch and yaw channels.
[0005] According to one aspect of the present invention, a method for electro-optical guidance and ship landing of an aircraft based on dual-channel coupling is provided, comprising the following eight steps: Step S10: The ship's electro-optical guidance system uses an infrared thermal imager to measure the aircraft's pitch and azimuth signals; the ship's electro-optical guidance system uses a laser rangefinder to measure the distance between the aircraft and the landing point; and a rate gyroscope is used to measure the aircraft's pitch rate signal, denoted as... The yaw rate signal of the aircraft is measured and recorded as... The approximate vertical deviation signal of the aircraft is calculated based on the distance information between the aircraft and the landing point and the aircraft's pitch signal; the approximate lateral deviation signal of the aircraft is calculated based on the distance information between the aircraft and the landing point and the aircraft's azimuth signal as follows: ; ; in To measure the pitch signal of the aircraft using an infrared thermal imager with an electro-optical guidance system, To measure the azimuth signal of the aircraft using an infrared thermal imager with an electro-optical guidance system, To measure the distance between the aircraft and the landing point using the laser rangefinder of the ship's electro-optical guidance system, This is the approximate vertical deviation signal for the aircraft; This is the approximate lateral deviation signal of the aircraft.
[0006] Step S20: Based on the pitch signal of the aircraft, design a leading inertial differential mixer to obtain a pitch inertial differential mixed signal; then integrate the pitch signal of the aircraft to obtain the aircraft pitch integral signal; finally, mix the vertical approximate deviation signal, pitch angular rate signal, aircraft pitch signal, aircraft pitch integral signal, and pitch inertial differential mixed signal to obtain the aircraft pitch composite signal.
[0007] Step S30: Based on the azimuth signal of the aircraft, design a leading inertial differential mixer to obtain an azimuth inertial differential mixed signal; then integrate the azimuth signal of the aircraft to obtain the aircraft azimuth integral signal; finally, mix the lateral approximate deviation signal, azimuth angular rate signal, aircraft azimuth signal, aircraft azimuth integral signal, and azimuth inertial differential mixed signal to obtain the comprehensive azimuth signal of the aircraft.
[0008] Step S40: Based on the pitch synthesis signal, design the pitch angle adaptive law, pitch angular velocity adaptive law, and pitch inertial differential hybrid adaptive law, and obtain the pitch angle adaptive coefficient, pitch angular velocity adaptive coefficient, and pitch inertial differential hybrid adaptive coefficient by integration respectively; then superimpose them to obtain the pitch direct interference compensation signal; then linearly superimpose the aircraft's pitch signal, vertical approximate deviation signal, pitch angular rate signal, and pitch inertial differential hybrid signal to obtain the pitch linear control signal.
[0009] Step S50: Based on the aforementioned azimuth composite signal, design the azimuth angle adaptive law, azimuth angular velocity adaptive law, and azimuth inertial differential hybrid adaptive law, and obtain the azimuth angle adaptive coefficient, azimuth angular velocity adaptive coefficient, and azimuth inertial differential hybrid adaptive coefficient by integration; then superimpose them to obtain the azimuth direct interference compensation signal; then linearly superimpose the aircraft's azimuth signal, vertical approximate deviation signal, azimuth angular rate signal, and azimuth inertial differential hybrid signal to obtain the azimuth linear control signal.
[0010] Step S60: Based on the pitch synthesis signal, design the azimuth hinge angle adaptive law, the azimuth hinge angular velocity adaptive law, and the azimuth hinge inertial differential hybrid adaptive law, and obtain the azimuth hinge angle adaptive coefficient, azimuth hinge angular velocity adaptive coefficient, and azimuth hinge inertial differential hybrid adaptive coefficient by integration, respectively; then superimpose them to obtain the pitch hinge interference compensation signal; finally, combine the azimuth inertial differential hybrid signal, the aircraft lateral approximate deviation signal, the yaw rate signal, and the aircraft azimuth signal to obtain the pitch hinge synthesis signal.
[0011] Step S70: Based on the aforementioned azimuth composite signal, design the pitch hinge angle adaptive law, pitch hinge angular velocity adaptive law, and pitch hinge inertial differential hybrid adaptive law, and obtain the pitch hinge angle adaptive coefficient, pitch hinge angular velocity adaptive coefficient, and pitch hinge inertial differential hybrid adaptive coefficient by integration, respectively; then superimpose them to obtain the azimuth hinge interference compensation signal; finally, combine the pitch inertial differential hybrid signal, the aircraft vertical approximate deviation signal, the pitch angular rate signal, and the aircraft pitch signal to obtain the azimuth hinge composite signal.
[0012] Step S80: The pitch linear control signal, pitch hinge composite signal, pitch direct interference compensation signal, yaw hinge interference compensation signal, and the aircraft's composite pitch signal are superimposed and combined to obtain the final aircraft electro-optical guidance pitch angle command signal, which is then sent to the aircraft pitch attitude tracking system for tracking, achieving vertical stability for the aircraft's electro-optical landing. Similarly, the azimuth linear control signal, azimuth hinge composite signal, azimuth direct interference compensation signal, pitch hinge interference compensation signal, and the aircraft's composite azimuth signal are superimposed and combined to obtain the final aircraft electro-optical guidance yaw angle command signal, which is then sent to the aircraft yaw attitude tracking system for tracking, achieving lateral stability for the aircraft's electro-optical landing.
[0013] In one exemplary embodiment of the present invention, based on the pitch signal of the aircraft, a leading inertial differential mixer is designed to obtain a pitch inertial differential mixed signal; then, the pitch signal of the aircraft is integrated to obtain the aircraft pitch integral signal; finally, the vertical approximate deviation signal, the pitch angular rate signal, the aircraft pitch signal, the aircraft pitch integral signal, and the pitch inertial differential mixed signal are mixed to obtain the aircraft's comprehensive pitch signal, including: ; ; ; in The pitch inertial differential hybrid signal is described above. , , For constant parameter signals, see the case implementation below for detailed selection. This is the pitch integral signal for the aircraft. For the aircraft's pitch synthesis signal, For constant parameter signals, see the case implementation below for detailed selection.
[0014] In one exemplary embodiment of the present invention, based on the azimuth signal of the aircraft, a leading inertial differential mixer is designed to obtain an azimuth-inertial differential mixed signal; then, the azimuth signal of the aircraft is integrated to obtain an integrated azimuth signal; finally, the azimuth comprehensive signal of the aircraft is obtained by mixing the lateral approximate deviation signal, the azimuth angular rate signal, the azimuth signal of the aircraft, the integrated azimuth signal of the aircraft, and the azimuth-inertial differential mixed signal, including: ; ; ; in The aforementioned azimuth inertial differential hybrid signal. This is the aircraft's azimuth integral signal. This is the overall orientation signal for the aircraft. For constant parameter signals, see the case implementation below for detailed selection.
[0015] In one exemplary embodiment of the present invention, based on the aforementioned pitch synthesis signal, a pitch angle adaptive law, a pitch angular velocity adaptive law, and a pitch inertial differential hybrid adaptive law are designed, and the pitch angle adaptive coefficient, pitch angular velocity adaptive coefficient, and pitch inertial differential hybrid adaptive coefficient are obtained by integration, respectively; these are then superimposed to obtain a pitch direct interference compensation signal; and the pitch signal, vertical approximate deviation signal, pitch angular rate signal, and pitch inertial differential hybrid signal of the aircraft are then linearly superimposed to obtain a pitch linear control signal including: ; ; ; ; in For pitch angle adaptive law, For pitch angular velocity adaptive law, It is a pitch inertial differential hybrid adaptive law. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients. This is the pitch angle adaptive coefficient. This is the pitch angular velocity adaptive coefficient. For pitch inertial differential hybrid adaptive coefficients. This is a signal used to compensate for direct pitch interference. For constant parameter signals, see the case implementation below for detailed selection. This is the pitch linear control signal.
[0016] In one exemplary embodiment of the present invention, based on the aforementioned azimuth composite signal, an azimuth angle adaptive law, an azimuth angular velocity adaptive law, and an azimuth inertial differential hybrid adaptive law are designed, and the azimuth angle adaptive coefficient, azimuth angular velocity adaptive coefficient, and azimuth inertial differential hybrid adaptive coefficient are obtained by integration, respectively; these are then superimposed to obtain the azimuth direct interference compensation signal; and the azimuth signal, vertical approximate deviation signal, azimuth angular rate signal, and azimuth inertial differential hybrid signal are then linearly superimposed to obtain the azimuth linear control signal, including: ; ; ; ; ; ; in This is the azimuth angle adaptive law. This is the adaptive law for azimuth angular velocity. It is a hybrid adaptive law of orientation inertial differential. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients. This is the azimuth angle adaptive coefficient. This is the adaptive coefficient for azimuth angular velocity. For the azimuth inertial differential hybrid adaptive coefficient. This is a direct interference compensation signal for azimuth. For constant parameter signals, see the case implementation below for detailed selection. This is the azimuth linear control signal.
[0017] In one exemplary embodiment of the present invention, based on the aforementioned azimuth composite signal, pitch hinge angle adaptive law, pitch hinge angular velocity adaptive law, and pitch hinge inertial differential hybrid adaptive law are designed, and the pitch hinge angle adaptive coefficient, pitch hinge angular velocity adaptive coefficient, and pitch hinge inertial differential hybrid adaptive coefficient are obtained by integration, respectively; then, they are superimposed to obtain the azimuth hinge interference compensation signal; finally, the azimuth hinge composite signal is obtained by combining the pitch inertial differential hybrid signal, the aircraft vertical approximate deviation signal, the pitch angular rate signal, and the aircraft pitch signal, including: ; ; ; in This is the adaptive coefficient for the pitch hinge angle; This is the adaptive coefficient for the pitch hinge angular velocity; For pitch hinge inertial differential hybrid adaptive coefficients. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients. This is the pitch hinge angle adaptive coefficient. For the pitch hinge angular velocity adaptive coefficient, For pitch hinge inertial differential hybrid adaptive coefficients. This is the azimuth hinge interference compensation signal. This is the integrated signal for the orientation hinge. , , , For constant parameter signals, see the case implementation below for detailed selection.
[0018] In one exemplary embodiment of the present invention, based on the pitch synthesis signal, an azimuth hinge angle adaptive law, an azimuth hinge angular velocity adaptive law, and an azimuth hinge inertial differential hybrid adaptive law are designed, and the azimuth hinge angle adaptive coefficient, azimuth hinge angular velocity adaptive coefficient, and azimuth hinge inertial differential hybrid adaptive coefficient are obtained by integration, respectively; then these are superimposed to obtain the pitch hinge interference compensation signal; finally, the pitch hinge synthesis signal is obtained by combining the azimuth inertial differential hybrid signal, the aircraft lateral approximate deviation signal, the yaw rate signal, and the aircraft azimuth signal, including: ; ; ; ; ; ; in This is the adaptive coefficient for the azimuth hinge angle; This is the adaptive coefficient for the angular velocity of the azimuth hinge; For the azimuth hinge inertial differential hybrid adaptive coefficient. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients. This is the adaptive coefficient for the azimuth hinge angle. This is the adaptive coefficient for the angular velocity of the azimuth hinge. For the azimuth hinge inertial differential hybrid adaptive coefficient. This is a pitch hinge interference compensation signal. This is the integrated signal for the pitch hinge. , , , For constant parameter signals, see the case implementation below for detailed selection.
[0019] In one exemplary embodiment of the present invention, the pitch linear control signal, pitch hinge composite signal, pitch direct interference compensation signal, yaw hinge interference compensation signal, and the aircraft's composite pitch signal are superimposed and combined to form the final aircraft electro-optical guidance pitch angle command signal; the azimuth linear control signal, azimuth hinge composite signal, azimuth direct interference compensation signal, pitch hinge interference compensation signal, and the aircraft's composite azimuth signal are superimposed and combined to form the final aircraft electro-optical guidance yaw angle command signal, including: ; ; in This is the pitch angle command signal for the aircraft's electro-optical guidance. This is a constant parameter; detailed selection can be found in the case study below. This is the yaw angle command signal for the aircraft's electro-optical guidance. This is a constant parameter; detailed selection can be found in the case study below.
[0020] Beneficial effects This invention discloses an electro-optical guidance landing method for aircraft based on dual-channel coupling. Its main innovations are as follows: First, it employs an adaptive method to classify, estimate, and compensate for the hinge coupling of the two channels. Second, it combines linear feedback control with hinge feedback control, effectively eliminating the disturbances and low accuracy issues caused by the dual-channel coupling of the aircraft in the landing control. This method is particularly suitable for the landing control of aircraft with strong nonlinearity and severe dual-channel coupling, giving this invention high engineering practical value.
[0021] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Attached Figure Description
[0022] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.
[0023] Figure 1 This is a flowchart of a flight vehicle photoelectric guidance landing method based on dual-channel coupling provided by the present invention.
[0024] Figure 2 This is the pitch signal curve (unit: degrees) of the aircraft provided by the method in the embodiments of the present invention. Figure 3 This is the azimuth signal curve (unit: degrees) of the aircraft provided by the method in the embodiments of the present invention. Figure 4 This is the distance information (unit: meters) between the aircraft and the landing point provided in the embodiments of the present invention. Figure 5 This is the aircraft pitch integral signal curve (unitless) of the method provided in the embodiments of the present invention. Figure 6 This is the aircraft azimuth integral signal curve (unitless) of the method provided in the embodiments of the present invention. Figure 7 This is the aircraft pitch linear control signal curve (unitless) provided by the method in the embodiments of the present invention. Figure 8 This is the aircraft azimuth linear control signal curve (unitless) of the method provided in the embodiments of the present invention. Figure 9 This is the aircraft altitude change curve (unit: meters) provided by the method in the embodiment of the present invention. Figure 10This is a curve showing the lateral position change of the aircraft according to the method provided in this embodiment of the invention (unit: meters). Detailed Implementation
[0025] Exemplary embodiments will now be described more fully with reference to the accompanying drawings. However, these exemplary embodiments can be implemented in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided to make the invention more comprehensive and complete, and to fully convey the concept of the exemplary embodiments to those skilled in the art. The described features, structures, or characteristics can be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to give a full understanding of embodiments of the invention. However, those skilled in the art will recognize that the technical solutions of the invention may be practiced with one or more of these specific details omitted, or other methods, components, apparatus, steps, etc., may be employed. In other instances, well-known technical solutions are not shown or described in detail to avoid obscuring various aspects of the invention.
[0026] This invention provides a ship landing guidance method for electro-optical guided aircraft using a dual-channel coupling approach. It measures the pitch azimuth and distance information of the aircraft using an infrared thermal imager and a laser rangefinder within the electro-optical guidance system, and calculates the approximate vertical and lateral deviation signals. Simultaneously, it obtains a comprehensive pitch azimuth signal by combining inertial differential, integral, and yaw rate data. An adaptive law is then designed to compensate for direct interference between the pitch and azimuth channels, and a linear control signal for pitch azimuth is constructed. Furthermore, an adaptive method is used to design a pitch azimuth hinge interference compensation signal to address the coupling between the two channels, and the comprehensive hinge feedback information for pitch azimuth is integrated, thus achieving electro-optical guided landing of the aircraft on a ship.
[0027] The following will, with reference to the accompanying drawings, further explain and illustrate a dual-channel coupled electro-optical guidance landing method for aircraft according to the present invention. (Reference) Figure 1 As shown, this method for electro-optical guidance and ship landing of an aircraft based on dual-channel coupling may include the following steps: Step S10: The ship's electro-optical guidance system uses an infrared thermal imager to measure the aircraft's pitch and azimuth signals; the ship's electro-optical guidance system uses a laser rangefinder to measure the distance between the aircraft and the landing point; and a rate gyroscope is used to measure the aircraft's pitch rate signal, denoted as... The yaw rate signal of the aircraft is measured and recorded as... The approximate vertical deviation signal of the aircraft is calculated based on the distance information between the aircraft and the landing point and the aircraft's pitch signal; the approximate lateral deviation signal of the aircraft is calculated based on the distance information between the aircraft and the landing point and the aircraft's azimuth signal as follows: ; ; To measure the pitch signal of the aircraft using an infrared thermal imager with an electro-optical guidance system, To measure the azimuth signal of the aircraft using an infrared thermal imager with an electro-optical guidance system, To measure the distance between the aircraft and the landing point using the laser rangefinder of the ship's electro-optical guidance system, This is the approximate vertical deviation signal for the aircraft; This is an approximate lateral deviation signal for the aircraft. Step S20: Based on the pitch signal of the aircraft, design a leading inertial differential mixer to obtain a pitch inertial differential mixed signal; then integrate the pitch signal of the aircraft to obtain the aircraft pitch integral signal; finally, mix the vertical approximate deviation signal, pitch angular rate signal, aircraft pitch signal, aircraft pitch integral signal, and pitch inertial differential mixed signal to obtain the aircraft pitch composite signal.
[0028] Specifically, based on the pitch signal of the aircraft, a lead inertial differential mixer is designed to obtain the pitch inertial differential mixed signal as follows: ; in The pitch inertial differential hybrid signal is described above. , , For constant parameter signals, see the case implementation below for detailed selection.
[0029] Secondly, the aircraft's pitch signal is integrated to obtain the aircraft's integrated pitch signal as follows: ; in This is the pitch integral signal for the aircraft.
[0030] Finally, based on the vertical approximate deviation signal, pitch rate signal, aircraft pitch signal, aircraft pitch integral signal, and pitch inertial differential mixed signal, the aircraft's comprehensive pitch signal is obtained as follows: ; in For the aircraft's pitch synthesis signal, For constant parameter signals, see the case implementation below for detailed selection.
[0031] Step S30: Based on the azimuth signal of the aircraft, design a leading inertial differential mixer to obtain an azimuth inertial differential mixed signal; then integrate the azimuth signal of the aircraft to obtain the aircraft azimuth integral signal; finally, mix the lateral approximate deviation signal, azimuth angular rate signal, aircraft azimuth signal, aircraft azimuth integral signal, and azimuth inertial differential mixed signal to obtain the comprehensive azimuth signal of the aircraft.
[0032] Specifically, based on the azimuth signal of the aforementioned aircraft, a lead-inertial differential mixer is designed to obtain the azimuth-inertial differential mixed signal as follows: ; in The aforementioned azimuth inertial differential hybrid signal.
[0033] Secondly, the aircraft's azimuth signal is integrated to obtain the aircraft's integrated azimuth signal as follows: ; in This is the aircraft's azimuth integral signal.
[0034] Finally, the comprehensive azimuth signal of the aircraft is obtained by mixing the vertical approximate deviation signal, the azimuth rate signal, the aircraft's azimuth signal, the aircraft's azimuth integral signal, and the azimuth inertial differential mixed signal as follows: ; in This is the overall orientation signal for the aircraft. For constant parameter signals, see the case implementation below for detailed selection.
[0035] Step S40: Based on the pitch synthesis signal, design the pitch angle adaptive law, pitch angular velocity adaptive law, and pitch inertial differential hybrid adaptive law, and obtain the pitch angle adaptive coefficient, pitch angular velocity adaptive coefficient, and pitch inertial differential hybrid adaptive coefficient by integration respectively; then superimpose them to obtain the pitch direct interference compensation signal; then linearly superimpose the aircraft's pitch signal, vertical approximate deviation signal, pitch angular rate signal, and pitch inertial differential hybrid signal to obtain the pitch linear control signal.
[0036] Specifically, based on the aforementioned pitch synthesis signal, the following pitch angle adaptive law, pitch angular velocity adaptive law, and pitch inertial differential hybrid adaptive law are designed: ; in For pitch angle adaptive law, For pitch angular velocity adaptive law, It is a pitch inertial differential hybrid adaptive law. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients.
[0037] Secondly, by integrating the above adaptive law, we obtain the pitch angle adaptive coefficient, pitch angular velocity adaptive coefficient, and pitch inertia differential hybrid adaptive coefficient as follows: ; in This is the pitch angle adaptive coefficient. This is the pitch angular velocity adaptive coefficient. For pitch inertial differential hybrid adaptive coefficients.
[0038] Then, the above adaptive coefficients are combined and superimposed to obtain the pitch direct interference compensation signal as follows: ; in This is a signal used to compensate for direct pitch interference.
[0039] Finally, the pitch signal, vertical approximate deviation signal, pitch angular rate signal, and pitch inertial differential mixed signal of the aircraft are linearly superimposed to obtain the following pitch linear control signal: ; in For constant parameter signals, see the case implementation below for detailed selection. This is the pitch linear control signal.
[0040] Step S50: Based on the aforementioned azimuth composite signal, design the azimuth angle adaptive law, azimuth angular velocity adaptive law, and azimuth inertial differential hybrid adaptive law, and obtain the azimuth angle adaptive coefficient, azimuth angular velocity adaptive coefficient, and azimuth inertial differential hybrid adaptive coefficient by integration; then superimpose them to obtain the azimuth direct interference compensation signal; then linearly superimpose the aircraft's azimuth signal, vertical approximate deviation signal, azimuth angular rate signal, and azimuth inertial differential hybrid signal to obtain the azimuth linear control signal.
[0041] Specifically, based on the aforementioned azimuth composite signal, the following adaptive laws are designed: azimuth angle adaptive law, azimuth angular velocity adaptive law, and azimuth inertial differential hybrid adaptive law: ; in This is the azimuth angle adaptive law. This is the adaptive law for azimuth angular velocity. It is a hybrid adaptive law of orientation inertial differential. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients.
[0042] Secondly, by integrating the above adaptive law, we obtain the azimuth angle adaptive coefficient, azimuth angular velocity adaptive coefficient, and azimuth inertia differential hybrid adaptive coefficient as follows: ; ; ; in This is the azimuth angle adaptive coefficient. This is the adaptive coefficient for azimuth angular velocity. For the azimuth inertial differential hybrid adaptive coefficient.
[0043] Then, the adaptive coefficients are combined and superimposed to obtain the azimuth direct interference compensation signal as follows: ; in This is a direct interference compensation signal for azimuth.
[0044] Finally, the azimuth signal, lateral approximate deviation signal, azimuth angular rate signal, and azimuth inertial differential mixed signal are linearly superimposed to obtain the following azimuth linear control signal: ; in For constant parameter signals, see the case implementation below for detailed selection. This is the azimuth linear control signal.
[0045] Step S60: Based on the pitch composite signal, design the azimuth hinge angle adaptive law, the azimuth hinge angular velocity adaptive law, and the azimuth hinge inertial differential hybrid adaptive law, and obtain the azimuth hinge angle adaptive coefficient, azimuth hinge angular velocity adaptive coefficient, and azimuth hinge inertial differential hybrid adaptive coefficient by integration, respectively; then superimpose them to obtain the pitch hinge interference compensation signal; finally, combine the azimuth inertial differential hybrid signal, the aircraft lateral approximate deviation signal, the yaw rate signal, and the aircraft azimuth signal to obtain the pitch hinge composite signal.
[0046] Specifically, based on the aforementioned pitch synthesis signal, the following adaptive laws are designed: azimuth hinge angle adaptive law, azimuth hinge angular velocity adaptive law, and azimuth hinge inertial differential hybrid adaptive law: ; in This is the adaptive coefficient for the azimuth hinge angle; This is the adaptive coefficient for the angular velocity of the azimuth hinge; For the azimuth hinge inertial differential hybrid adaptive coefficient. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients.
[0047] Then, by integrating the above adaptive law, we obtain the azimuth hinge angle adaptive coefficient, azimuth hinge angular velocity adaptive coefficient, and azimuth hinge inertia differential hybrid adaptive coefficient as follows: ; ; ; in This is the adaptive coefficient for the azimuth hinge angle. This is the adaptive coefficient for the angular velocity of the azimuth hinge. For the azimuth hinge inertial differential hybrid adaptive coefficient.
[0048] Secondly, the pitch hinge interference compensation signal is obtained by superimposing the above adaptive coefficients as follows: ; in This is a pitch hinge interference compensation signal.
[0049] Finally, the pitch hinge composite signal is obtained by combining the azimuth inertial differential mixed signal, the aircraft lateral approximate deviation signal, the yaw rate signal, and the aircraft's azimuth signal: ; in This is the integrated signal for the pitch hinge. , , , For constant parameter signals, see the case implementation below for detailed selection.
[0050] Step S70: Based on the aforementioned azimuth composite signal, design the pitch hinge angle adaptive law, pitch hinge angular velocity adaptive law, and pitch hinge inertial differential hybrid adaptive law, and obtain the pitch hinge angle adaptive coefficient, pitch hinge angular velocity adaptive coefficient, and pitch hinge inertial differential hybrid adaptive coefficient by integration, respectively; then superimpose them to obtain the azimuth hinge interference compensation signal; finally, combine the pitch inertial differential hybrid signal, the aircraft vertical approximate deviation signal, the pitch angular rate signal, and the aircraft pitch signal to obtain the azimuth hinge composite signal.
[0051] Specifically, based on the aforementioned azimuth synthesis signal, the following adaptive laws are designed: pitch hinge angle adaptive law, pitch hinge angular velocity adaptive law, and pitch hinge inertial differential hybrid adaptive law: ; in This is the adaptive coefficient for the pitch hinge angle; This is the adaptive coefficient for the pitch hinge angular velocity; For pitch hinge inertial differential hybrid adaptive coefficients. , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients.
[0052] Then, by integrating the above adaptive law, we obtain the pitch hinge angle adaptive coefficient, pitch hinge angular velocity adaptive coefficient, and pitch hinge inertia differential hybrid adaptive coefficient as follows: in This is the pitch hinge angle adaptive coefficient. For the pitch hinge angular velocity adaptive coefficient, For pitch hinge inertial differential hybrid adaptive coefficients.
[0053] Secondly, the azimuth hinge interference compensation signal is obtained by superimposing the above adaptive coefficients as follows: ; in This is the azimuth hinge interference compensation signal.
[0054] Finally, the azimuth hinge composite signal is obtained by combining the pitch inertial differential hybrid signal, the aircraft's vertical approximate deviation signal, the pitch angular rate signal, and the aircraft's pitch signal: ; in This is the integrated signal for the orientation hinge. , , , For constant parameter signals, see the case implementation below for detailed selection.
[0055] Step S80: The pitch linear control signal, pitch hinge composite signal, pitch direct interference compensation signal, yaw hinge interference compensation signal, and the aircraft's composite pitch signal are superimposed and combined to obtain the final aircraft electro-optical guidance pitch angle command signal, which is then sent to the aircraft pitch attitude tracking system for tracking, achieving vertical stability for the aircraft's electro-optical landing. Similarly, the azimuth linear control signal, azimuth hinge composite signal, azimuth direct interference compensation signal, pitch hinge interference compensation signal, and the aircraft's composite azimuth signal are superimposed and combined to obtain the final aircraft electro-optical guidance yaw angle command signal, which is then sent to the aircraft yaw attitude tracking system for tracking, achieving lateral stability for the aircraft's electro-optical landing.
[0056] Specifically, the pitch linear control signal, pitch hinge composite signal, pitch direct interference compensation signal, yaw hinge interference compensation signal, and aircraft pitch composite signal are first superimposed and combined to obtain the final aircraft electro-optical guidance pitch angle command signal as follows: ; in This is the pitch angle command signal for the aircraft's electro-optical guidance. This is a constant parameter; detailed selection can be found in the case study below.
[0057] Secondly, by superimposing and combining the aforementioned azimuth linear control signal, azimuth hinge integrated signal, azimuth direct interference compensation signal, pitch hinge interference compensation signal, and the aircraft's azimuth integrated signal, the final aircraft electro-optical guidance yaw angle command signal is obtained as follows: ; in This is the yaw angle command signal for the aircraft's electro-optical guidance. This is a constant parameter; detailed selection can be found in the case study below.
[0058] Finally, the aforementioned aircraft electro-optical guidance pitch angle command signal and aircraft electro-optical guidance yaw angle command signal are respectively transmitted to the pitch angle tracking subsystem and yaw angle tracking subsystem of the aircraft control system to achieve attitude stabilization and tracking, thereby realizing the aircraft's electro-optical landing guidance. The pitch angle tracking subsystem and yaw angle tracking subsystem belong to the aircraft flight control system; however, this invention mainly focuses on electro-optical landing guidance, and therefore will not be elaborated upon further.
[0059] Case Implementation and Computer Simulation Results Analysis In step S10, the pitch signal of the aircraft is measured using an infrared thermal imager of the photoelectric guidance system, such as... Figure 2 As shown; then, the infrared thermal imager of the photoelectric guidance system is used to measure the azimuth signal of the aircraft, such as... Figure 3As shown. The distance between the aircraft and the landing point is measured using a laser rangefinder from the ship's electro-optical guidance system, such as... Figure 4 As shown.
[0060] In step S20, select , , , The aircraft pitch integral signal is obtained as follows: Figure 5 As shown.
[0061] In step S30, select The aircraft's azimuth integral signal is obtained as follows: Figure 6 As shown.
[0062] In step S40, , , , The pitch linear control signal is obtained as follows: Figure 7 As shown.
[0063] In step S50, , , , The azimuth linear control signal is obtained as follows: Figure 8 As shown.
[0064] In step S60, , , , , , , .
[0065] In step S70, , , , , , , .
[0066] In step S80, select ; The final aircraft altitude change curve is shown in Figure 9, and the final aircraft lateral position change curve is shown in Figure 10. (From...) Figure 2 It can be seen that the initial pitch angle is around 20 degrees; from Figure 3 It can be seen that the initial azimuth angle is around -9 degrees; from Figure 9 It can be seen that the initial height was 250 meters, from Figure 10It can be seen that the initial lateral deviation was -100 meters, and ultimately, under the guidance of photoelectric sensors, the aircraft accurately landed on the ship within 15 seconds. The overall experimental results show that the method provided by this invention exhibits good stability and accuracy, demonstrating the effectiveness of this method based on decoupling compensation photoelectric guidance with dual-channel coupling, and possessing high engineering application and promotion value.
Claims
1. A method for electro-optical guidance and ship landing of an aircraft based on dual-channel coupling, characterized in that, Includes the following steps: Step S10: The ship's electro-optical guidance system uses an infrared thermal imager to measure the aircraft's pitch and azimuth signals; the ship's electro-optical guidance system uses a laser rangefinder to measure the distance between the aircraft and the landing point; and a rate gyroscope is used to measure the aircraft's pitch rate signal, denoted as... The yaw rate signal of the aircraft is measured and recorded as... The approximate vertical deviation signal of the aircraft is calculated based on the distance information between the aircraft and the landing point and the aircraft's pitch signal; the approximate lateral deviation signal of the aircraft is calculated based on the distance information between the aircraft and the landing point and the aircraft's azimuth signal as follows: ； ； in To measure the pitch signal of the aircraft using an infrared thermal imager with an electro-optical guidance system, To measure the azimuth signal of the aircraft using an infrared thermal imager with an electro-optical guidance system, To measure the distance between the aircraft and the landing point using the laser rangefinder of the ship's electro-optical guidance system, This is the approximate vertical deviation signal for the aircraft; This is an approximate lateral deviation signal for the aircraft. Step S20: Based on the aircraft's pitch signal, a leading inertial differential mixer is designed to obtain a pitch inertial differential mixed signal; then, the aircraft's pitch signal is integrated to obtain the aircraft's pitch integral signal; finally, the vertical approximate deviation signal, pitch angular rate signal, aircraft pitch signal, aircraft pitch integral signal, and pitch inertial differential mixed signal are mixed to obtain the aircraft's comprehensive pitch signal as follows: ； ； ； in The pitch inertial differential hybrid signal is described above. , , For constant parameter signals; For aircraft pitch integration signal; For the aircraft's pitch synthesis signal, For constant parameter signals; Step S30: Based on the azimuth signal of the aircraft, design a leading inertial differential mixer to obtain an azimuth inertial differential mixed signal; then integrate the azimuth signal of the aircraft to obtain the aircraft azimuth integrated signal. The comprehensive azimuth signal of the aircraft is obtained by mixing the lateral approximate deviation signal, azimuth angular rate signal, aircraft azimuth signal, aircraft azimuth integral signal, and azimuth inertial differential mixed signal as follows: ； ； ； in This refers to the aforementioned azimuth inertial differential hybrid signal; This is the aircraft's azimuth integral signal; This is the overall orientation signal for the aircraft. For constant parameter signals; Step S40: Based on the aforementioned pitch synthesis signal, design the pitch angle adaptive law, pitch angular velocity adaptive law, and pitch inertial differential hybrid adaptive law, and obtain the pitch angle adaptive coefficient, pitch angular velocity adaptive coefficient, and pitch inertial differential hybrid adaptive coefficient respectively through integration; then superimpose them to obtain the pitch direct interference compensation signal; then linearly superimpose the aircraft's pitch signal, vertical approximate deviation signal, pitch angular rate signal, and pitch inertial differential hybrid signal to obtain the following pitch linear control signal: ； ； ； ； Among them For pitch angle adaptive law, For pitch angular velocity adaptive law, For pitch inertial differential hybrid adaptive law; , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients; This is the pitch angle adaptive coefficient. This is the pitch angular velocity adaptive coefficient. The pitch inertial differential hybrid adaptive coefficient; For pitch direct interference compensation signal; For constant parameter signals; This is the pitch linear control signal; Step S50: Based on the aforementioned azimuth composite signal, design the azimuth angle adaptive law, azimuth angular velocity adaptive law, and azimuth inertial differential hybrid adaptive law, and obtain the azimuth angle adaptive coefficient, azimuth angular velocity adaptive coefficient, and azimuth inertial differential hybrid adaptive coefficient respectively through integration; then superimpose them to obtain the azimuth direct interference compensation signal; then linearly superimpose the aircraft's azimuth signal, vertical approximate deviation signal, azimuth angular rate signal, and azimuth inertial differential hybrid signal to obtain the following azimuth linear control signal: ； ； ； ； ； ； in This is the azimuth angle adaptive law. This is the adaptive law for azimuth angular velocity. This is a hybrid adaptive law of azimuth inertial differential; , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients; This is the azimuth angle adaptive coefficient. This is the adaptive coefficient for azimuth angular velocity. For the azimuth inertial differential hybrid adaptive coefficient; This is a direct interference compensation signal for azimuth. For constant parameter signals; This is the azimuth linear control signal; Step S60: Based on the pitch synthesis signal, design the azimuth hinge angle adaptive law, the azimuth hinge angular velocity adaptive law, and the azimuth hinge inertial differential hybrid adaptive law, and obtain the azimuth hinge angle adaptive coefficient, azimuth hinge angular velocity adaptive coefficient, and azimuth hinge inertial differential hybrid adaptive coefficient by integration, respectively; then superimpose them to obtain the pitch hinge interference compensation signal; finally, combine the azimuth inertial differential hybrid signal, the aircraft lateral approximate deviation signal, the yaw rate signal, and the aircraft azimuth signal to obtain the pitch hinge synthesis signal as follows: ； ； ； ； ； ； in This is the adaptive coefficient for the azimuth hinge angle; This is the adaptive coefficient for the angular velocity of the azimuth hinge; For the azimuth hinge inertial differential hybrid adaptive coefficient; , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients; This is the adaptive coefficient for the azimuth hinge angle. This is the adaptive coefficient for the angular velocity of the azimuth hinge. For the azimuth hinge inertial differential hybrid adaptive coefficient; For pitch hinge interference compensation signal; For pitch hinge integrated signals; , , , For constant parameter signals; Step S70: Based on the aforementioned azimuth composite signal, design the pitch hinge angle adaptive law, pitch hinge angular velocity adaptive law, and pitch hinge inertial differential hybrid adaptive law, and obtain the pitch hinge angle adaptive coefficient, pitch hinge angular velocity adaptive coefficient, and pitch hinge inertial differential hybrid adaptive coefficient respectively through integration; then superimpose them to obtain the azimuth hinge interference compensation signal; finally, combine the pitch inertial differential hybrid signal, the aircraft vertical approximate deviation signal, the pitch angular rate signal, and the aircraft's pitch signal to obtain the following azimuth hinge composite signal: ； ； ； in This is the adaptive coefficient for the pitch hinge angle; This is the adaptive coefficient for the pitch hinge angular velocity; The pitch hinge inertial differential hybrid adaptive coefficient; , , This is a constant adaptive parameter used to adjust the convergence speed of the adaptive coefficients; This is the pitch hinge angle adaptive coefficient. For the pitch hinge angular velocity adaptive coefficient, The pitch hinge inertial differential hybrid adaptive coefficient; This is a signal for compensating for interference from the azimuth hinge. This is a composite signal for the orientation hinge; , , , For constant parameter signals; Step S80: The pitch linear control signal, pitch hinge composite signal, pitch direct interference compensation signal, yaw hinge interference compensation signal, and the aircraft's composite pitch signal are superimposed and combined to obtain the final aircraft electro-optical guidance pitch angle command signal, which is sent to the aircraft pitch attitude tracking system for tracking, thereby achieving vertical stability of the aircraft's electro-optical landing; The azimuth linear control signal, azimuth hinge composite signal, azimuth direct interference compensation signal, pitch hinge interference compensation signal, and the aircraft's composite azimuth signal are superimposed and combined to obtain the final aircraft electro-optical guidance yaw angle command signal, which is sent to the aircraft yaw attitude tracking system for tracking, thereby achieving lateral stability of the aircraft's electro-optical landing as follows: ； ； in This is the pitch angle command signal for the aircraft's electro-optical guidance. It is a constant parameter; This is the yaw angle command signal for the aircraft's electro-optical guidance. This is a constant parameter.