Helicopter post-survey aerial survey planning system guided by geophysical anomaly target area

Abstract

The present application relates to the field of aerogeophysical exploration and automatic flight control technology, and particularly relates to a helicopter posterior aerial survey planning system based on geophysical anomaly target area guidance. The present application generates a front space gradient tensor pulse, and innovatively constructs a geophysical morphological characteristic parameter to realize dynamic adjustment of flight speed and direction and accurate detection of geological anomaly areas, aiming at the problem that the fixed flight path in the traditional aerial survey method does not match the dynamic geological characteristics. The present application integrates aerodynamic control calculation equation sets to ensure safe and smooth flight, and introduces a flight attitude intensity quantitative evaluation mechanism to adjust the sampling frequency in real time and compensate for signal distortion caused by maneuvering. The present application effectively improves the data sampling density and spatial resolution, solves the problem that the fixed path in the traditional method does not match the geological characteristics and the data quality is affected by flight maneuvering, and improves the efficiency and accuracy of aerogeophysical exploration.

Description

Technical Field

[0001] This invention relates to the field of airborne geophysical exploration and automatic flight control technology, and more specifically, to a helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance. Background Technology

[0002] In airborne geophysical exploration, acquiring high-resolution, high-signal-to-noise-ratio data on the distribution of subsurface physical fields is crucial for identifying geological structures and mineral resources. To achieve this, the flight platform needs to accurately cover the pre-defined survey area and maintain a stable flight attitude to reduce measurement noise. Currently, the industry commonly uses airborne survey planning methods based on pre-defined gridded survey lines. This involves the aircraft flying at a constant speed and altitude along a series of equally spaced parallel or orthogonal grid lines, performing a uniform, carpet-like scan of the entire survey area. After data acquisition, post-processing is used to identify and delineate anomaly areas. However, existing technologies based on fixed-flight terrain blind scanning (such as referencing conventional airborne magnetic regulations and static surface map inversion techniques) have significant detection limitations. If key anomalies are detected using this approach, it may be necessary to plan a second or more intensive flight missions to obtain more detailed data.

[0003] However, this traditional approach based on fixed flight path planning has certain limitations. First, the method exhibits a lag in response to geological features; it cannot immediately adjust flight strategies upon detecting anomalies, potentially causing the aircraft to fly too quickly over critical anomaly zones with drastically changing gradients. This results in insufficient data sampling density, affecting the detailed characterization of the anomaly's morphology and boundaries. Second, the method of indiscriminately scanning the entire survey area consumes the same amount of flight resources in vast, geologically flat regions as in key anomaly zones, requiring improvement in exploration efficiency. Furthermore, while some existing anomaly tracking technologies can guide the aircraft to the anomaly center, their control logic is often relatively simplistic, focusing only on heading adjustments without deeply coupling and co-optimizing flight dynamics parameters such as flight speed and turn rate with the spatial morphological characteristics of the anomaly (such as its orientation and steepness).

[0004] Furthermore, the measurement accuracy of airborne geophysical sensors is extremely sensitive to changes in the attitude of the flight platform. When the aircraft performs maneuvers such as turning, acceleration, and deceleration, its own motion and structural vibrations generate electromagnetic interference, or the way the sensors cut through the Earth's magnetic field changes due to attitude variations, thus introducing significant noise into the measurement data, and even causing data distortion or temporary interruption. Existing technologies rarely consider the negative impact of flight maneuvers on data acquisition quality when planning flight paths, and lack a closed-loop control mechanism that can actively compensate for the data quality degradation caused by maneuvers while performing them to track geological anomalies.

[0005] Based on the above problems, this invention aims to solve the problems of mismatch between fixed flight paths and dynamic geological anomaly characteristics in traditional aerial surveying methods, as well as the lack of closed-loop coordinated control between aircraft maneuvering and data acquisition quality. Summary of the Invention

[0006] In view of this, in order to solve the problems mentioned in the background technology, a helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance is proposed.

[0007] The objective of this invention can be achieved through the following technical solution: This invention provides a helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance, comprising: a leading edge spatial gradient tensor pulse generation module, which acquires the physical field transient signal stream sampled by the original hardware layer collected by the helicopter airborne probe, and performs transient three-dimensional differential decomposition and singular value decomposition through a hardware calculation engine to generate a leading edge spatial gradient tensor pulse.

[0008] The geophysical morphological feature parameter generation module extracts the physical anomaly gradient magnitude vector and the main feature evolution direction from the frontal spatial gradient tensor pulse, and combines them to generate geophysical morphological feature parameters.

[0009] The equation generation module obtains the helicopter's current flight attitude package and extracts flight dynamics extreme boundary data. It then performs forced coverage and assimilation by combining geological and physical morphological parameters to generate a set of aerodynamic control equations.

[0010] The flight operation command sequence generation module inputs the update time constant of the current space physics exploration node into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, generate a geological morphology space-following flight operation command sequence and issue it.

[0011] The flight attitude severity quantification array generation module monitors the real-time mechanical response force sequence during the execution of the geological morphology space follow-up flight operation command sequence, and generates a real-time flight attitude severity quantification array by combining the aircraft composite maneuver attitude evaluation example model.

[0012] The clock limit latch instruction generation module uses a real-time flight attitude drastic measurement array to determine the instantaneous hardware detection signal distortion attenuation factor, generates an airborne probe limit clock frequency latch instruction, and achieves tomographic data restoration by forcibly writing back to trigger limit saturation clock sampling.

[0013] Compared with the prior art, the embodiments of the present invention have at least the following advantages or beneficial effects: (1) The present invention constructs geophysical morphological characteristic parameters and strongly correlates them with flight dynamics, thereby dynamically adjusting the flight speed and heading of the helicopter according to the real-time morphology of the underground anomaly. This mechanism uses the real-time calculated physical field spatial gradient magnitude as the reciprocal penalty factor for speed control and assimilates its main characteristic evolution direction as the main guiding source for heading control. This data-driven control method enables the helicopter to automatically decelerate when approaching or passing through areas with severe physical field gradients and to fly precisely along the direction of the anomaly's extension, while in areas with gentle geological backgrounds, it resumes normal cruising speed. Thus, intelligent on-demand allocation of detection resources is realized, and the data sampling density and spatial resolution of key geological anomaly target areas are improved without the need for manual intervention or secondary flights.

[0014] (2) This invention establishes a set of aerodynamic control equations that take into account the limits of flight dynamics, enabling the flight platform to maintain attitude stability and operational safety while performing geological morphology-guided flight. Before generating flight commands, this method pre-retrieves performance boundary data such as the maximum tilt angle and minimum yaw thrust available in the current state of the helicopter and incorporates these constraints into the solution process of the control equations. This mechanism ensures that all calculated maneuver control parameters are within the safety envelope of the aircraft. Ultimately, it achieves a balance between the autonomy of geological target tracking and the inherent aerodynamic performance limitations of the aircraft, reducing the risk of flight attitude instability caused by unreasonable commands generated by the autonomous system.

[0015] (3) This invention introduces a real-time quantitative assessment of the severity of flight attitude and injects it back into the front-end detection hardware to dynamically adjust the sampling frequency, thus constructing a closed-loop feedback adjustment path between flight control and data acquisition quality. This method can simultaneously monitor the physical deformation overhead of the helicopter's transmission mechanical components when executing maneuver commands and quantitatively assess the amplitude of unplanned flutter of the fuselage. Once the assessed severity of attitude exceeds a certain threshold, the system forcibly increases the sampling frequency of the geophysical sensor through a hardware interrupt channel. This mechanism can proactively predict and compensate for the distortion or interruption of the detection signal caused by flight maneuvering. Ultimately, it achieves forced filling and supplementation of detection data at the moment when the flight attitude is most unstable, eliminating data acquisition blind spots caused by the autonomous maneuvering of the aircraft and improving the continuity and reliability of the final geophysical data results. Attached Figure Description

[0016] To more clearly illustrate the technical solutions of the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0017] Figure 1 This is a schematic diagram of the system module structure connection of the present invention. Detailed Implementation

[0018] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0019] Please see Figure 1 This invention provides a helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance, including: a leading edge spatial gradient tensor pulse generation module, a geophysical morphological feature parameter generation module, a solution equation generation module, a flight operation command sequence generation module, a flight attitude severity quantification array generation module, and a clock limit latch command generation module.

[0020] The frontier spatial gradient tensor pulse generation module is connected to the geophysical morphological feature parameter generation module, the geophysical morphological feature parameter generation module is connected to the equation set generation module, the equation set generation module is connected to the flight operation command sequence generation module, the flight operation command sequence generation module is connected to the flight attitude severity quantification array generation module, and the flight attitude severity quantification array generation module is connected to the clock limit latch command generation module.

[0021] The leading-edge spatial gradient tensor pulse generation module acquires the transient physical field signal stream sampled by the original hardware layer from the helicopter's airborne probe, and generates leading-edge spatial gradient tensor pulses by performing transient three-dimensional differential decomposition and singular value decomposition through the hardware solution engine.

[0022] In a specific embodiment of the present invention, the transient signal stream of the physical field sampled by the original hardware layer collected by the helicopter airborne probe is obtained, and transient three-dimensional differential calculation and singular value decomposition are performed by the hardware calculation engine to generate the leading edge spatial gradient tensor pulse, including: importing the transient signal stream of the physical field sampled by the original hardware layer to the hardware calculation engine to perform transient three-dimensional differential calculation and extracting transient feature parameters of the physical field.

[0023] In a specific embodiment of the present invention, the transient physical field signal stream sampled by the original hardware layer is imported into the hardware calculation engine to perform transient three-dimensional differential calculation and extract transient physical field feature parameters, including: starting the airborne multi-probe induction coil and optical pump magnetometer equipment distributed in a three-dimensional spatial orthogonal array to enter the continuous sampling mode, and obtaining the transient physical field signal stream sampled by the original hardware layer of the continuous waveform in the basic target area.

[0024] By using high-pass filtering or moving average background removal operations, the slowly changing static Earth's main magnetic field and regional field background signals are effectively filtered out from the transient signal stream of the physical field sampled from the original hardware layer.

[0025] The high-frequency disturbances caused by underground geological anomalies are retained, and the transient characteristic parameters of the physical field are extracted.

[0026] Specifically, this embodiment further details the process of obtaining the periodic pitch compensation and tail rotor yaw angle compensation. After receiving the continuous dynamic state target calculated by the front-end aerodynamic control equations, the integrated flight control and detection processing unit needs to convert it into discrete physical compensation quantities that can be directly received and accurately executed by the mechanical actuators (servo drive module). In the longitudinal dynamic channel, the system first seamlessly intercepts the real-time flow data of the helicopter, including inertial sensing results, through the internal flight control bus interface, and extracts the absolute ground speed vector component of the aircraft at the current moment to obtain the current actual speed. Subsequently, the system calculates the transient difference between the target forward speed (the command speed forcibly constrained by the physical field gradient magnitude) derived from the aerodynamic control fundamental equations and the current actual speed. Considering that the change in the helicopter's forward speed is heavily dependent on the longitudinal tilt angle of the main rotor disk's total potential surface (i.e., longitudinal periodic pitch), and that the aerodynamic response surface of the main rotor generating translational thrust exhibits extremely strong nonlinear characteristics at different reference airspeeds, simple PID proportional feedback is difficult to achieve accurate and overshoot-free follow-up. Therefore, based on the speed difference calculated above, and combined with the current actual speed as a two-dimensional index pointer, the system consults the helicopter inverse dynamics model data table pre-stored in non-volatile memory. Through table lookup comparison and high-frequency surface interpolation algorithm addressing, the system directly extracts the pitch-back yaw degree of the main rotor thrust surface necessary to compensate for the speed error (such as preventing severe deceleration from overshooting), thereby accurately establishing the periodic pitch compensation amount. Simultaneously, on the yaw-controlled path in the track plane, the system extracts the target yaw rate generated in real-time by solving the aerodynamic equations, pointing towards the geologically anomalous, strongly extended structural surface. To convert the continuously distributed angular velocity in the time domain into a static angular displacement preset amount suitable for hard node control, the system introduces the inherent beat timing of this control closed loop—that is, the current space physics detection node update time constant (i.e., the equivalent time step of the detection sensor sampling and calculation cycle). Finally, considering that when the helicopter enters a servo yaw turn, the tail rotor not only needs to overcome the time-varying anti-torque caused by the dynamic flapping of the main rotor, but also needs to provide additional yaw acceleration to approximate the target yaw rate. The system takes the target yaw rate calculated from the equations, the actual yaw rate transmitted back from the flight control, and the update time constant of the current space physics detection node as joint inputs, calls the tail rotor aerodynamic performance inversion function, calculates the tail rotor thrust difference that must be compensated to establish the target yaw rate, and then, through thrust pitch mapping lookup table, rigorously calculates the collective pitch angle reference that the tail rotor hub mechanical actuator needs to be turned, thereby strictly obtaining the tail rotor yaw angle compensation degree.

[0027] In the above formula, The calculated longitudinal velocity difference of the helicopter is represented by units of meters per second. ; The current target speed after processing with meteorological and anti-runaway constraints. The current actual speed of the returning machine is represented by both units. The second formula describes the process of solving the longitudinal inverse dynamics problem. The periodic pitch compensation amount represents the output, and its unit physical dimension corresponds to radians. ; This represents the parameter matrix surface corresponding to the preset helicopter inverse dynamics model data table. This represents the two-dimensional nonlinear geometric interpolation function called by the hardware processor at the lower level, based on the two input velocity parameters mentioned above. The third formula illustrates the dynamic aerodynamic mapping of the tail rotor angular offset. The tail rotor yaw angle compensation degree calculated under the current cycle time is defined. The system uses the target yaw rate. Compared with the actual yaw rate The difference divided by the instruction duration The required yaw acceleration is obtained and combined with the actual yaw rate as the addressing key. This is then used to query the preset tail rotor inverse dynamics model data table. The two-dimensional nonlinear surface is used to extract the preset amount of the tail rotor blade's angle of attack deflection intervention, which fully meets the dynamic aerodynamic feedback trim requirements of the control level.

[0028] The difference between the target's forward velocity and the current actual velocity reflects the net longitudinal maneuvering requirement the helicopter needs to establish within the extremely short control refresh cycle to match the steep gradient of the target area. A reversal surge in the difference usually indicates that the aircraft is about to fly over a drastically changing gravitational or magnetic anomaly peak and urgently needs to brake and remain airborne. The preset helicopter inverse dynamics model data table is essentially a trim and transient control mapping matrix surface obtained by the aircraft manufacturer in wind tunnel or flight test calibration. It instantly establishes a rapid mapping between "motion requirement" and "rotor pitch mechanical servo offset" across complex aerodynamic differential equations to ensure "zero delay" in response. The periodic pitch compensation specifically refers to the tilt angle increment command applied to the longitudinal normal of the main rotor's automatic swashplate. The target yaw rate is the horizontal rotation speed of the fuselage required to enable the front-end probe to accurately perform follow-up scanning along the anomaly path. The tail rotor yaw angle compensation is a parameter that reflects the total change in the angle of attack required by each blade in the tail rotor hub to overcome air resistance and achieve the target angular rate.

[0029] For example, the system listens and determines that the current space physics probe node update time constant (i.e., the probe / resolve update cycle) is set to... When cruising to the boundary of an abnormally high gradient region, the solution of the fundamental aerodynamic control equations is strongly suppressed by the gradient magnitude, requiring the output safe target forward velocity to be... The current actual speed captured by the bus is The system immediately calculates the longitudinal velocity difference: The system indicates a strong need to decelerate (to counteract the forward momentum). The system then sets the current reference airspeed ( ) and speed demand reduction ( As , Using the combined index key value, the built-in helicopter inverse dynamics model data table is consulted. Through nearest neighbor address lookup and bilinear surface interpolation, it is determined that to achieve this level of damped deceleration at this speed, the longitudinal cyclic pitch mechanism needs to provide a rearward thrust equivalent to... Precise trim can only be achieved with a propeller disk pitch-back aerodynamic rudder effect of a certain degree. Therefore, the cyclic pitch compensation amount needs to be extracted and output. On the other hand, to adapt to the abnormal spatial characteristics, the target yaw rate calculated in real time by the equations is for a right turn. The system calculates the angular acceleration demand based on the difference between the target yaw rate and the real-time yaw rate, and uses this value as the simultaneous addressing key to consult the tail rotor inverse dynamics model data table for inversion, thus calculating the... Ultimately, the system simultaneously calculates the longitudinal rotor micro-management correction amount used for "vehicle speed control" ( ) and tail rotor collective pitch fine-tuning for "leadership tracking" ( This means that the periodic pitch compensation and tail rotor yaw angle compensation are successfully obtained, making the subsequent parallel frame packet delivery active, verifiable, and highly transparent.

[0030] Calculate the partial derivatives of the transient characteristic parameters of the physical field in the orthogonal triaxial directions and assemble them into a three-dimensional spatial gradient matrix.

[0031] Singular value decomposition is performed on the three-dimensional spatial gradient matrix to obtain the leading spatial gradient tensor impulse.

[0032] Specifically, this embodiment discloses a helicopter-based a posteriori aerial survey planning method guided by geophysical anomaly target areas. The main execution unit of the method is a flight control and detection integrated processing unit deployed within the helicopter. The initial step of the method is that the system first sends initialization commands to the airborne three-dimensional array of induction coils and optical pump magnetometers, activating the aforementioned detection devices to enter a continuous high-frequency sampling mode, thereby continuously capturing the transient signal stream of the underlying physical field in the basic target area space. This transient signal stream of the underlying physical field is a series of time-series vector data carrying three-axis components. Subsequently, the hardware calculation engine within the flight control and detection integrated processing unit, such as a field-programmable gate array (FPGA), receives the transient signal stream of the underlying physical field in real time. The hardware calculation engine executes a transient three-dimensional differential calculation program, filtering out slowly changing static background signals such as the Earth's main magnetic field and regional fields from the original signal through high-pass filtering or moving average background removal operations, retaining only the high-frequency disturbance part caused by underground geological anomalies, and then extracting the transient characteristic parameters of the physical field. To quantify the drastic spatial changes in the anomalous field, the spatial partial derivatives of the transient characteristic parameters of the physical field along the orthogonal three axes in the helicopter flight path coordinate system need to be calculated. This calculation process combines real-time velocity vector information provided by the onboard inertial navigation unit (IMU) and uses a chain rule to convert the rate of change of time into the rate of change of space. The processing unit assembles the nine calculated partial derivative values ​​into a three-dimensional spatial gradient matrix of three dimensions to characterize the omnidirectional change of the anomalous field at the discrete spatial sampling point where the current probe is located. This calculation process abandons the "mathematically indeterminate and unsolvable" error of using a single probe to convert spatial gradients through a chain rule of velocity and time change. Instead, it utilizes the synchronous and pure signals collected by a three-dimensional orthogonal probe array consisting of the helicopter's nose, stern strut, and wingtips. Using the known specific physical distance between the probes in a rigid system as the denominator, direct spatial difference physics operations are performed to accurately reconstruct mutually independent and dimensionally consistent ( The system uses nine independent spatial partial derivatives of the anomalous field. Finally, in order to analyze the primary and secondary directions and intensity of the change, the system calls the singular value decomposition (SVD) algorithm to numerically solve the three-dimensional spatial gradient matrix. The resulting singular value matrix and the left and right singular vector matrices together constitute the final output of this step, which is a leading-edge spatial gradient tensor pulse that reflects the steepest local change direction, secondary change direction and corresponding change amplitude of the anomalous field.

[0033] In the above formula This represents the gradient matrix in three-dimensional space. , , These represent the components of the transient characteristic parameters of the physical field along the orthogonal triaxial directions, respectively, with units of Tesla. Or Tesla . , , Represents spatial coordinates, with the unit being meters. Therefore, the matrix The dimensions of each element in the medium are: . and These are the left and right singular vector matrices obtained through singular value decomposition. They are all orthogonal matrices and dimensionless. It is a diagonal matrix with singular values ​​on its diagonal, and the dimensions of its diagonal elements are the same as those of the matrix. The elements have the same dimensions, and are also of the same type. This ensures that the formula is self-consistent in terms of physical dimensions.

[0034] The underlying physical field transient signal stream refers to the raw magnetic field strength time series data collected by an optical pump magnetometer at a specific sampling frequency. This sampling frequency is an adjustable parameter determined based on the expected scale of the target geological body and the helicopter's cruising speed, typically set between 200Hz and 1kHz to ensure sufficient analytical capability for fine geological structures. The hardware processing engine specifically refers to a field-programmable gate array (FPGA) or digital signal processor (DSP) capable of nanosecond-level logic operations and parallel processing. This type of hardware is chosen to meet the millisecond-level real-time feedback required by flight control. The physical field transient characteristic parameters are residual signals obtained by stripping the low-frequency static environment background from the underlying physical field transient signal stream. They primarily reflect the rapid magnetic field fluctuations caused by the aircraft traversing local geological anomalies. The leading-edge spatial gradient tensor pulse is a set of data structures obtained by singular value decomposition of the three-dimensional spatial gradient matrix, containing a diagonal matrix. Its diagonal elements are singular values. and two orthogonal matrices and This data pulse is generated once at each detection node, providing real-time input for subsequent flight attitude calculation.

[0035] For example, suppose in At a given moment, the optical pump magnetometer inside the helicopter acquires a frame of transient physical field signal stream reflecting the local magnetic field environment at a sampling frequency of 500Hz. After filtering out the static background, the time derivative value of the extracted transient characteristic parameter of the physical field is: Meanwhile, the real-time velocity vector of the helicopter transmitted back by the airborne inertial navigation unit is... The hardware calculation engine within the integrated flight control and detection processing unit, based on the aforementioned data, assembles the three-dimensional spatial gradient matrix of the current sampling point through approximate calculations. The specific values ​​are: the first row is The second line is The third line is All units are Subsequently, the system immediately processed the three-dimensional spatial gradient matrix. Performing singular value decomposition yields a frontal spatial gradient tensor pulse. This pulse specifically comprises a singular value diagonal matrix. Its diagonal elements are Left singular vector matrix transpose of the right singular vector matrix These are two specific numerical representations of third-order orthogonal matrices. This set consists of matrices... , and The resulting frontal spatial gradient tensor pulse is then used as the final output of this step and sent to the next processing stage.

[0036] The geophysical morphological feature parameter generation module extracts the physical anomaly gradient magnitude vector and the main feature evolution direction from the frontal spatial gradient tensor pulse, and combines them to generate geophysical morphological feature parameters.

[0037] In a specific embodiment of the present invention, the physical anomaly gradient magnitude vector and the main feature evolution direction are extracted from the frontal spatial gradient tensor pulse and combined to generate geophysical morphological feature parameters, including: calculating the sum of squares of all diagonal elements of the singular value diagonal matrix in the frontal spatial gradient tensor pulse and performing a square root operation to obtain the total absolute variation as the physical anomaly gradient magnitude vector.

[0038] The feature space is analyzed and compared with the leading spatial gradient tensor impulse, and a set of orthogonal feature vectors corresponding to the maximum eigenvalues ​​are extracted as the main feature evolution direction.

[0039] By combining the gradient magnitude vector of physical anomalies with the evolution direction of the main features, geological and physical morphological feature parameters are generated.

[0040] Specifically, after acquiring the leading-edge spatial gradient tensor pulse, the flight control and detection integrated processing unit immediately enters the calculation process for this step. The unit first internally retrieves the singular value diagonal matrix contained in the leading-edge spatial gradient tensor pulse. The diagonal elements of this matrix represent the degree of variation of the physical field along the three orthogonal principal directions. To obtain a single scalar measure of the overall strength of the anomaly, the system performs a sum of squares on all diagonal elements of this singular value diagonal matrix and takes the square root of the result. This operation is equivalent to calculating the Frobenius norm of the original spatial gradient matrix, and the numerical result is taken as the total absolute variability at the current 3D spatial node and defined as a scalar vector of the physical anomaly gradient magnitude. Simultaneously, to determine the dominant extension direction of the anomaly in space, the system performs a feature space alignment on the leading spatial gradient tensor impulses. This feature space analytical alignment specifically refers to the alignment of the three singular values ​​contained in the diagonal matrix. , , The values ​​are compared one by one and sorted in descending order to accurately locate the singular value of the largest value. The process involves, specifically, extracting the maximum singular value from the left and right singular vector matrices that constitute the tensor after numerical comparison. The corresponding left singular eigenvector This vector points in the direction of the steepest gradient in the physical field space (i.e., perpendicular to the direction of anomaly extension). To enable the helicopter to fly in accordance with the edge of the anomaly, the system projects this feature vector onto the horizontal plane and rotates it 90 degrees counterclockwise, extracting its orthogonal geometric normal direction as the main feature evolution direction representing the geological structure's strike. Finally, the system combines the scalar form of the physical anomaly gradient magnitude vector calculated in the previous step with the currently extracted vector form of the main feature evolution direction through data structure packaging to generate the output of this step. This output is a geophysical morphological feature parameter that simultaneously possesses both an instantaneous steepness indicator of the physical field and a quantitative indicator of the strike spatial extension, used to directly and inversely constrain the helicopter's flight dynamics model.

[0041] In the above formula, This represents the gradient magnitude vector of physical anomalies, i.e., the three-dimensional spatial gradient matrix. Frobenius norm . , , These are the three singular values ​​contained in the leading-edge spatial gradient tensor impulse, with units of . .therefore, The dimensions are also It has the same dimensions as the physical field gradient, and is dimensionally consistent. The direction of evolution of the main features is represented by the maximum singular value. The corresponding left singular vector . It is a unit vector, dimensionless, and represents a purely spatial direction.

[0042] The physical anomaly gradient magnitude vector is a scalar value that quantifies the overall intensity of changes in the physical field at the current location. A larger value indicates a clearer boundary or greater difference in physical properties of the geological anomaly the helicopter is traversing. The main characteristic evolution direction is a three-dimensional unit vector indicating the direction of the greatest gradient change in the geological anomaly in three-dimensional space; this typically corresponds directly to the strike of a geological structure or the extension of a ore body. The geophysical morphological characteristic parameters are a composite data structure, usually organized as a quadruple containing a scalar and a three-dimensional vector. It integrates the anomaly's intensity and direction information, forming the physical basis for direct intervention in the flight control system.

[0043] For example, continuing from the calculation results in step S1, the leading-edge spatial gradient tensor pulse received by the flight control and detection integrated processing unit contains a singular value diagonal matrix. Its three singular values ​​on the diagonal are respectively , , The system first calculates the physical anomaly gradient magnitude vector, then substitutes the numerical values ​​into the calculation: ,get The value is approximately Simultaneously, the system resolves the left singular vector corresponding to the maximum singular value of 21.01 from the leading edge spatial gradient tensor impulses. Assume the specific value of this vector obtained in the aforementioned S1 operation is... This vector is then determined as the direction of evolution of the backbone feature. Ultimately, the system will use a scalar... with vector Combining these parameters generates geophysical morphological characteristics, whose data structure is represented as follows: This is then used as output to provide the next step of aerodynamic control solution.

[0044] The equation generation module obtains the helicopter's current flight attitude package and extracts flight dynamics extreme boundary data. It then performs forced coverage and assimilation by combining geological and physical morphological parameters to generate a set of aerodynamic control equations.

[0045] In a specific embodiment of the present invention, the current flight attitude package of the helicopter is obtained and the flight dynamics extreme boundary data is extracted. Combined with the geological and physical morphological feature parameters, the data is forcibly covered and assimilated to generate a set of aerodynamic control basic solution equations. This includes: reading the current flight attitude package of the helicopter and extracting the maximum tilt angle limit of the rotor and the minimum available yaw thrust of the tail rotor from the physical extreme value envelope table built into the helicopter flight control system, and combining them into flight dynamics extreme boundary data.

[0046] The physical anomaly gradient magnitude vector attached to the geological and physical morphological features is extracted as the reciprocal penalty factor to force the overlay of the basic forward velocity scalar mapping formula.

[0047] The evolution direction of the main features in the geophysical morphological parameters is forcibly assimilated into the main guiding input source of the tail rotor aerodynamic yaw rate. Combined with the basic forward velocity scalar mapping formula that has been covered and changed and the extreme boundary data of flight dynamics, the aerodynamic control basic solution equation set is obtained by integration and recombination.

[0048] Specifically, after generating the geophysical morphological parameters, the integrated flight control and detection processing unit immediately initiates the calculation for this step. The system uses its internal data bus to read the attitude data packets transmitted back from the helicopter flight control computer in real time via high-frequency polling, and extracts key flight dynamics constraint parameters, including the maximum allowable forward tilt angle limit of the rotor shaft and the minimum effective yaw thrust value that the tail rotor can provide under the current flight conditions. These parameters are combined and packaged to form a localized flight dynamics extreme boundary data, serving as a hard constraint for subsequent control calculations. Next, the processing unit precisely extracts the scalar component from the geophysical morphological parameters generated in the previous step, namely the physical anomaly gradient magnitude vector. This vector is used as a reciprocal penalty factor, applied in multiplication or division to forcibly override the preset basic forward velocity scalar mapping formula within the helicopter's independent automatic cruise control center, meaning that the planned velocity is inversely proportional to the physical field gradient intensity. Simultaneously, the processing unit continues to extract the vector component, i.e., the evolution direction of the main features, from the geophysical morphological characteristic parameters and forcibly assimilates it, making it the main guiding input source for the helicopter tail rotor aerodynamic yaw rate. This means that the extension direction of the geological anomaly will directly dominate the aircraft's heading rate of change. Finally, the system integrates the new basic forward velocity scalar mapping formula obtained after the reciprocal penalty factor coverage change, the yaw control law guided by the aforementioned evolution direction of the main features, and the flight dynamics extreme boundary data. These three are recombined to form the core output of this embodiment, namely a strongly coupled aerodynamic control fundamental solution equation set driven by geophysical characteristics.

[0049] The above formulas form the core of the fundamental aerodynamic control equations. In the first formula, This is a newly generated forward speed command, and its unit is... . This is the basic forward speed calibration value provided by the automatic cruise module, in units of... . It is a dimensionless velocity damping coefficient with a dimension of 1, used to adjust the magnitude vector of the physical anomaly gradient. The intensity of the speed suppression. The magnitude of the physical anomaly gradient is calculated in real time, and its unit is... ; The system's preset baseline gradient magnitude constant, whose unit is also 1. This is used to ensure that the product of coefficients in the denominator achieves dimensionless consistency. Based on this, a velocity range constraint is introduced, where... This represents the maximum speed constraint corresponding to the maximum forward tilt angle of the aircraft. This represents the lower bound protection threshold of the system's preset minimum safe and effective cruise speed for helicopters, preventing the aircraft from entering hovering loss of control or vortex ring state due to extreme gradients. It ensures the system's feasibility and dimensional consistency. This speed command is affected by the maximum tilt angle. Corresponding maximum speed The constraints ensure its executability and dimensional consistency. In the second and third formulas, This is the final output yaw rate command, in units of... . It is the yaw control proportional gain, in units of . It is the difference between the target heading and the current heading, in radians. The target heading evolves from the main characteristics. horizontal components and The solution is obtained by using the arctangent function, thus ensuring the uniqueness of the course calculation across all quadrants. This is the current real-time heading angle read from the flight control bus, in units of... . It is a saturation function, representing the minimum physical yaw rate limit that can be converted from the tail rotor thrust limit. and the maximum physical yaw rate limit The active truncation constraint on the yaw rate ensures the self-consistency of dimensions and physical constraints.

[0050] The flight dynamics extreme boundary data is a dynamically updated dataset containing the mechanical performance limits achievable by the helicopter rotor and tail rotor under current flight conditions. These parameters are calculated by the flight control main computer based on real-time data such as air pressure, temperature, and load. The reciprocal penalty factor is a control logic implementation method. Its core idea is that the greater the detected geophysical field gradient, the lower the flight speed is suppressed, enabling precise scanning of anomalous areas. The basic forward velocity scalar mapping formula is the benchmark calculation formula used by the automatic cruise system to maintain a constant cruise speed when there is no geological signal guidance. The main guidance input source represents the main characteristic evolution direction and has the highest priority at the control level; its generated yaw commands will override regular waypoint navigation commands. The aerodynamic control fundamental solution equations are a series of coupled mathematical expressions that directly transform geophysical measurement results into flight dynamics control variables. Velocity damping coefficient. This is a key preset parameter, whose value is determined to ensure that the helicopter can decelerate smoothly without excessive pitch turbulence in areas of rapid gradient change. It is usually set through flight simulation or actual test flights. to Within the range.

[0051] For example, the system inherits the geophysical morphological feature parameters generated in the previous embodiment. The integrated flight control and detection processing unit obtains the current flight dynamics extreme boundary data—the maximum permissible forward speed—through high-frequency polling of the data bus. for The maximum yaw rate is The system sets the speed damping coefficient. for The basic forward speed given by the automatic cruise module for The system first calculates the speed command and substitutes it into the formula: This value is less than The limit is reached, therefore it is valid. Next, the yaw command is calculated, and the horizontal component of the main feature evolution direction is extracted from the geophysical morphological parameters. Assuming the current helicopter heading... For the East, that is The target heading is Then the heading difference The minimum turning angle obtained after winding is Set the yaw control gain. for The calculated yaw rate command is Because this value exceeds The limit of the saturation function Once effective, the final output yaw rate command is limited to... Therefore, the solution to the integrated aerodynamic control fundamental equation set at this moment is: the target forward velocity. Target yaw rate (Turn right).

[0052] The flight operation command sequence generation module inputs the update time constant of the current space physics exploration node into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, generate a geological morphology space-following flight operation command sequence and issue it.

[0053] In a specific embodiment of the present invention, the current space physics exploration node update time constant is input into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, generate a sequence of geological morphology space follow-up flight operation commands and issue them, including: inputting the current space physics exploration node update time constant into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, and obtaining the periodic pitch compensation amount and the tail rotor yaw angle compensation degree.

[0054] In a specific embodiment of the present invention, the current space physics detection node update time constant is input into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, and the periodic pitch compensation amount and tail rotor yaw angle compensation degree are obtained. This includes: based on the difference between the target forward speed calculated by the aerodynamic control fundamental solution equation set and the current actual speed, consulting the preset helicopter inverse dynamics model data table, and calculating the periodic pitch compensation amount.

[0055] The required angular acceleration is calculated by using the difference between the target yaw rate and the current actual yaw rate obtained from the aerodynamic control fundamental equations. Combined with the pre-set helicopter tail rotor aerodynamic efficiency inversion function mapping table, the tail rotor yaw angle compensation degree is obtained.

[0056] Specifically, this embodiment further details the process of restoring tomographic data by forcibly writing back to trigger the limit saturation clock sampling. After the flight control and detection integrated processing unit detects that the flight attitude has exceeded the limit and generates a sampling surge command, in order to avoid the inherent queuing delay and protocol stack parsing time of conventional airborne data buses (such as CAN or Ethernet buses) in multi-node communication, the system chooses to directly skip the software application layer and activate an extremely low-latency peripheral dedicated interrupt response channel via the underlying hardware pin. In terms of hardware architecture, the peripheral dedicated interrupt response channel is directly physically bridged from the general purpose input / output (GPIO) high-speed hardware interrupt pin of the field programmable gate array (FPGA) in the flight control and detection integrated processing unit to the external clock input terminal of the trigger logic control board inside the optical pump magnetometer through a shielded anti-interference cable. Through this physical direct connection channel, the system forcibly writes back the packaged airborne probe limit clock frequency latch command in the form of nanosecond level level toggling to the front-end controlled hardware—that is, directly erases and writes the original trigger timing gate array (FPGA control register) that controls the pulse period of the optical pump magnetometer. This preemptive write-back action, which transcends the conventional handshake protocol, instantly relinquishes control of the uniform, low-frequency sampling timing originally maintained by the magnetometer's internal reference timer with the highest hardware priority. When a helicopter makes a rapid, steep turn or ascends based on geophysical anomalies, the linear and angular velocities of the mounted detection pod undergo a nonlinear superposition and surge. Maintaining the original sampling rate would drastically increase the physical distance between adjacent effective detection points, leaving unidentifiable "data gaps" or signal ambiguity at critical anomaly boundaries. Therefore, this write-back mechanism forcibly triggers the physical sensing probe to overcome conventional energy-saving and data redundancy limitations during a moment of high spatial overload. The system continuously acquires data at the physical saturation upper limit frequency, activating the extreme saturation clock sampling mode. By "flooding" massive amounts of high-density discrete magnetic field slice data into the storage unit within a few-second maneuver window, the system successfully recovers the truncated fault data that would inevitably be missed under conventional sampling. This allows for the objective restoration and maintenance of the spatial topological coherence of the finely structured anomaly target area in subsequent 3D mesh inversion.

[0057] In the above formulas, the first formula describes the hard handover model of sampling timing control. The sampling frequency that the probe is actually excited at the current moment, expressed in Hertz. or ; This is the system's default normal sampling frequency. The full-frequency sampling rate is the inherent physical limit of the probe's induction coil or optical pump; both have dimensions of 1. ; This is a Boolean step function triggered by a dedicated interrupt response channel; it is dimensionless and guarantees the self-consistency of the equation's dimensions. The second formula defines the process for constructing the captured dataset. Represents the truncated tomographic window ( arrive A high-density discrete dataset sequence that is forcibly filled in. Let be the Dirac impulse function, used for simulating time-continuous fields. Discrete points are extracted using a new, extremely high sampling frequency. The third formula establishes the physical constraints for topological coherence. This represents the actual physical displacement step size of two consecutive sampling points in three-dimensional space, in meters. Molecular part The probe sensor (pod) velocity during helicopter translation was calculated using rigid body kinematics formulas. With high angular velocity The magnitude of the sum of absolute instantaneous linear velocities under superposition, in units of The division of linear velocity and frequency precisely restores the distance dimension. ; The maximum permissible spatial fault threshold to ensure that the topological boundaries of the geological anomaly target area are not distorted, in units of This inequality mathematically proves that at the numerator end where the velocity of motion increases dramatically, the sampling frequency at the denominator end must be increased exponentially in order to maintain the spatial resolution below the threshold.

[0058] In one specific embodiment of the present invention, The maximum permissible spatial fault threshold, measured in meters (m), is the threshold value allowed to ensure that the topological boundaries of the geological anomaly target area are not distorted. In practical high-precision airborne geophysical exploration projects, this threshold is typically set at [value missing]. to Between, the typical working value is set as Its design is primarily based on the strong coupling mechanism between the "spatial Nyquist sampling theorem" and the "minimum subdivision accuracy of the posterior 3D inversion mesh": Since the target tracked by the system will excite a high-frequency abrupt change physical field with an extremely short wavelength and a very steep gradient in the spatial domain, in order to prevent "spatial domain signal aliasing" and the smoothing and blurring of topological boundaries by the algorithm during the post-processing 3D data mesh reconstruction, the spatial Euclidean distance between adjacent effective discrete detection points must be strictly less than half of the minimum effective spatial wavelength of the anomalous field; limiting it to the centimeter level ensures, from both theoretical and engineering perspectives, that even if the helicopter's external probe is thrown out to an extremely high absolute linear velocity during maneuvering with the long extension rod, the system can still offset the long displacement by surging the hardware sampling frequency, and comprehensively scan the extremely clear real topological geometric edges of the complex underground target body with a dense, non-uniform "sub-meter-level micro-fault slice", thereby effectively avoiding microscopic coherence distortion caused by missed sampling during maneuvering.

[0059] Among them, the dedicated interrupt response channel for peripherals specifically refers to an independent hard-wired I / O channel distinct from the system's main communication bus, such as a dedicated GPIO high-speed interrupt pin or parallel bus. It has nanosecond-level response characteristics and can block CPU preemption by all other non-fatal processes. The airborne probe's limit clock frequency latch instruction is a sequence of machine code used to directly overwrite the underlying configuration of the sensor's ADC (analog-to-digital converter) and clock divider matrix. The original trigger timing gate array is a core logic device (such as a CPLD or FPGA) installed inside the optical pump magnetometer to distribute pulse gating period signals according to a preset division ratio. The momentary large-overload torsional moment in space refers to a brief physical state where the helicopter, based on the abnormal target area's physical field strength characteristics, forcibly calculates and executes pitch or yaw evasive maneuvers, causing the fuselage and external pylons to experience a surge in acceleration and a sudden change in attitude angle. Limit saturation clock sampling refers to a working mode where the sampling rate of the detector's optical polarization gate or sensing oscillation circuit is briefly pushed to the device's physical saturation limit without considering long-term heat dissipation and data storage capacity pressure. The spatial topological coherence of the target area refers to the fact that when forming geophysical maps, the edge contours, morphological connections, and contour lines of geological anomalies (such as vein strikes) do not suffer from artificial distortion, breakage, or artifacts due to the sparse sampling points caused by the movement process.

[0060] For example, the listening module detected that the helicopter was... Cruising at ground speed, and guided by the physics field gradient tensor, instantly cutting into a position with an angular velocity of up to [missing information]. (about The helicopter makes a sharp turn to meet the abnormal boundary of the qualified bureau. It is known that the optical pump magnetometer pod is installed at a distance of [distance missing] from the helicopter's center of gravity. On the fuselage extension rod. At the moment of turning, the absolute spatial linear velocity experienced by the probe surges to approximately [value missing]. The system's original conventional sampling frequency was... If no intervention is taken at this point, the distance between two adjacent detection points in space will be stretched to... Because the helicopter was under a state of high overload with severe vibrations ( This can easily cause the optical pump signal to momentarily "defocus" or briefly lose lock, leading to a disruption in this already sparse area. The presence of numerous blank and invalid values ​​in the step size data directly disrupts the fault continuity threshold (set threshold). At this moment, the integrated flight control and detection processing unit immediately writes the high-energy latch instruction at address 0x0F back to the probe's FPGA via the dedicated peripheral interrupt response channel, suspending its original timing array and instantly entering... The extreme saturation clock sampling mode. In this mode, the new spatial sampling step size is forced to be compressed to... Much smaller than the threshold requirement Within a 0.8-second large torsional maneuver window, the system injects high-density, extremely high-density transient data points in a "parallel transient" manner. Even if 20% of these data points are filtered out and deemed invalid due to vibration noise, the remaining 3200 effective measurement points can still reproduce the abrupt change in magnetic field intensity at the twisted fault location in a good and coherent manner, much like a macro scan. This effectively restores the spatial topological coherence of the target area of ​​the complex ore body and achieves a reverse closed-loop compensation where "the more intense the maneuver, the more refined the local exploration."

[0061] Based on the periodic pitch compensation, a forced anti-runaway deceleration command is generated, and based on the tail rotor yaw angle compensation, a tail rotor heading follow-up yaw operation intervention command is generated.

[0062] The forced anti-runaway deceleration command and the tail rotor heading follow-up deflection operation intervention command are encoded with overlapping bit mask and merged into a sequence of geological morphology space follow-up flight operation commands for issuance.

[0063] Specifically, after constructing the fundamental aerodynamic control equations in the preceding steps, the integrated flight control and detection processing unit immediately begins solving these equations and generating executable commands. The processing unit first introduces a system-level time base parameter, namely the current space physics detection node update time constant, which is synchronized with the physics field sampling period in S1. Next, the processing unit uses this time constant as the iteration step size to perform an online numerical iterative convergence solution to the fundamental aerodynamic control equations. This solution process aims to calculate the specific mechanical aerodynamic control compensation amount required to bring the helicopter's flight state closer to the target defined by the equations within the next time step. Specifically, based on the difference between the calculated target forward velocity and the current actual velocity, the system calculates the periodic pitch compensation amount required to reduce the rotor longitudinal tilt angle to achieve this velocity change by consulting a pre-set helicopter inverse dynamics model data table. Simultaneously, the system multiplies the calculated target yaw rate by the current space physics detection node update time constant to obtain the tail rotor yaw angle compensation degree required to achieve heading tracking. After acquiring these two core compensation values, the system generates two types of logical commands based on them: First, a high-priority forced anti-exit deceleration command is generated based on the periodic pitch compensation value; second, a tail rotor yaw angle compensation command is generated based on the tail rotor yaw angle compensation value. Finally, to ensure the efficiency and atomicity of command transmission, the system performs overlapping bitmasking encoding on the forced anti-exit deceleration command and the tail rotor yaw angle compensation command according to the flight control control allocation rules defined in the airborne underlying communication protocol. The two are then merged into a unified data frame containing the complete flight intent, namely the geological morphology space-following flight operation command sequence, and transmitted to the mechanical actuators of the main rotor and tail rotor via the internal high-speed data bus.

[0064] The first formula above describes the periodic pitch compensation amount. The calculation method uses degrees or radians as its unit. It is achieved through an inverse dynamic surface interpolation function, based on the current velocity read from the flight control system. And the longitudinal velocity difference derived from the aerodynamic control fundamental equations The result is obtained from a lookup table. The second formula calculates the tail rotor yaw angle compensation. The unit is degrees or radians. It is the target yaw rate calculated using the aerodynamic control fundamental equations. Combined with the current yaw rate The difference between the two is divided by the angular acceleration requirement obtained from the probe update time constant, and then jointly inverted using the tail rotor trim mapping table. The third formula illustrates the overlapping bit mask encoding process. It is the final generated sequence of geological morphology space-following flight operation commands (usually a 16-bit or 32-bit integer). This represents a quantization function that converts floating-point compensation values ​​into fixed-bit integer instruction codes. This represents a bitwise left shift operation. Representing a bitwise OR operation, this formula encodes the longitudinal and yaw commands separately and then packages them into the same command word, with the dimension being a unitless numeric code.

[0065] In a specific embodiment of the present invention, the overlapping bit mask encoding here refers to a digital packing technique that, under the condition of limited underlying bus bandwidth, assigns non-overlapping register bit segments (such as the high 8 bits and the low 8 bits) to different control parameters, and then uses logical left shift to assign addresses and logical "OR" operation to merge them.

[0066] The current space physics detection node update time constant is a time interval bound to the probe sampling frequency in S1, typically within... Up to 20 Between these parameters, it defines the response rate of the entire closed-loop control system. Cyclic pitch compensation refers to the angle value of the main rotor's longitudinal cyclic pitch change required to achieve the target speed. Tail rotor yaw angle compensation refers to the angle value of the fuselage collective pitch that the tail rotor blades need to change to track the target heading rate of change. Forced anti-runaway deceleration commands and tail rotor heading follow-up yaw operation intervention commands are the digital and protocol-based representations of the above compensation quantities; they are the final electronic signals sent to the servo controller. Overlapping bitmask encoding is a data compression and packetization technique that allows multiple independent control commands to be sent simultaneously within a single data transmission cycle. Different commands are distinguished by predefined data bit allocation, thereby improving the efficiency of the communication bus and ensuring control synchronization.

[0067] For example, using the calculation result of S3, the target's forward speed is The target yaw rate is The system obtains the current actual forward speed from the helicopter's flight control unit. The system's current space physics detection node update time constant is set to... or First, the system calculates the periodic pitch compensation amount. This is based on the inverse dynamics model. To reduce speed from Reduce to It is necessary to apply a The longitudinal cyclic pitch change is 0 degrees. Therefore, the cyclic pitch compensation is -0.5 degrees. Next, the tail rotor yaw angle compensation is calculated: , approximately equal to The system then quantizes and encodes these two compensation values. Assume the protocol specifies a 16-bit word command, with the high 8 bits for pitch compensation and the low 8 bits for yaw compensation, and quantization precisions of 0.01 degrees / bit and 0.005 degrees / bit respectively. A pitch compensation of -0.5 degrees is quantized to -50 and encoded as follows: (Binary two's complement). A yaw compensation of 0.149 degrees is quantized to $29.8, rounded to 30, and encoded as follows: Perform overlapping bitmask encoding; the instruction word is... Ultimately, the system generates a value containing hexadecimal values. The sequence of flight operation commands based on the geological morphology space is transmitted to the corresponding servo controller via the CAN bus.

[0068] The flight attitude severity quantification array generation module monitors the real-time mechanical response force sequence during the execution of the geological morphology space follow-up flight operation command sequence, and generates a real-time flight attitude severity quantification array by combining the aircraft composite maneuver attitude evaluation example model.

[0069] In a specific embodiment of the present invention, the real-time mechanical response force sequence during the execution of the geological morphology space-following flight operation command sequence is monitored, and combined with the aircraft compound maneuver attitude evaluation case model, a real-time flight attitude severity quantification array is generated, including: cyclically monitoring the tension and friction feedback waveforms returned by the sensors on the main rotor periodic pitch control rod and the tail rotor drive servo connection shaft of the helicopter, and obtaining the real-time mechanical response force sequence within the set time-series rolling window.

[0070] The cumulative rate of dynamic change in yaw servo angular velocity and the absolute motion delay offset time of each mechanical operating surface are accumulated within the force sequence of the real-time mechanical response.

[0071] By substituting the cumulative rate of dynamic change in yaw servo angular velocity and the absolute action delay offset time into the aircraft's high-maneuver attitude evaluation case model, a real-time flight attitude severity quantification array is generated.

[0072] Specifically, the integrated flight control and detection processing unit immediately initiates the listening and processing tasks of this step simultaneously when issuing the sequence of geological morphology spatial follow-up flight operation commands. The processing unit cyclically monitors the force sensor connected to the helicopter main rotor periodic pitch control rod and the torque sensor coupled to the tail rotor drive servo connection shaft, acquiring the returned tension and friction feedback waveforms at high frequency. These analog signals, after analog-to-digital conversion, are cached within a set time-series rolling window, collectively forming a multi-channel real-time mechanical response force sequence. This sequence records in detail all the instantaneous overhead of physical deformation of the actuator during the response to flight commands. Subsequently, the processing unit processes the cached data within this sequence. First, it extracts the angular velocity data returned by the tail rotor servo motor encoder, calculates its first-order difference within the time window, and accumulates its absolute value to obtain the cumulative rate of dynamic change of yaw servo angular velocity, which is used to quantify the smoothness of yaw control. Secondly, the system compares the timestamps of the geological morphology spatial servo flight operation command sequence with the moments when the sensor readings of each mechanical operating surface in the real-time mechanical response force sequence first show a significant deviation from the static baseline, calculating the absolute action delay offset time for each channel. Finally, the processing unit takes the calculated cumulative rate of dynamic change in yaw servo angular velocity and the absolute action delay offset time as input, and feeds them in parallel into a pre-stored aircraft high-maneuver attitude evaluation example model. This model, through weighted summation, generates a real-time flight attitude severity quantification array that can quantitatively characterize the unplanned flutter amplitude generated by the fuselage structure during servo flight, and uses this as the final output of this step.

[0073] The first formula above defines the cumulative rate of dynamic change in yaw servo angular velocity. In essence, the length should be set to Within the measurement time window, yaw servo angular acceleration The time integral average of the absolute value, in units of The formula follows the standard mean-value integral theorem, dividing the integral value by the complete window sampling duration to ensure the mathematical accuracy of the transient impact energy distribution evaluation. The second formula calculates the... Absolute motion delay offset time of each mechanical operating surface It is the response time. With the time the command was issued The difference is expressed in seconds. The third formula is the core expression of the aircraft's high-maneuver attitude assessment model, used to calculate the total attitude severity score. . and These are preset dimensionless weighting coefficients. To eliminate the influence of different physical quantities, the formula uses... and Each used its own normalization constant. and Standardization was performed to ensure the final score. It is a dimensionless value, and the dimensions are correct.

[0074] In one specific embodiment of the present invention, and These are preset dimensionless weighting coefficients. In actual airborne geophysical exploration projects, The typical value for (normalized angular acceleration cumulative rate weight) is usually set to 0.6, and the sum of the delay weights for each mechanical operating surface is... The typical value is set to 0.4. This weighting ratio is based on the principle of helicopter aeroelastic coupling and the noise sensitivity mechanism of optical magnetometers: during follow-up flight tracking geological anomalies, the dynamic changes in angular velocity caused by active maneuver commands (corresponding to...) This is the primary and most direct source of high-frequency "whipping" elastic vibration directly causing the helicopter's long tail boom and external geophysical pod to vibrate. This high-frequency vibration can easily and instantly break down the phase-locked frequency of the optical pump, leading to severe data distortion or loss; while the action delay of the mechanical operating surface (corresponding to...) The primary effects are phase lag in collective pitch control and low-frequency swaying of the fuselage, with relatively mild transient destructive energy to sensor geomagnetic cutting noise. Therefore, based on a large amount of vibration sensing data from actual flight tests and a multivariate regression analysis model of magnetic field baseline noise, the active yaw acceleration, which directly induces high-frequency oscillations, is weighted at 0.6, and the mechanical response hysteresis, which induces low-frequency swaying, is weighted at 0.4. This model can most objectively and accurately fit the nonlinear destructive effect of the aircraft's actual physical attitude deterioration under compound maneuvers on the quality of geophysical data.

[0075] The real-time mechanical response force sequence is a data matrix where rows represent time sampling points and columns represent different physical sensor channels, such as the Z-axis tension of the main torque converter rod and the XY-plane torque of the tail rotor servo axis. The length of the timing scroll window is typically set to... Up to 500 This length is sufficient to capture the entire response process of a maneuver while ensuring real-time calculation. The cumulative rate of dynamic change in yaw servo angular velocity is an indicator of whether control commands are too abrupt; a high cumulative rate implies frequent tail rotor jerks, potentially causing fuselage vibration. The absolute action delay offset time reflects the response performance of the mechanical transmission system and is a key parameter for measuring whether the control system can keep up with commands. The aircraft's high-maneuver attitude evaluation simulation model is essentially a multi-input, single-output linear weighted model, with weighting coefficients... and The system was calibrated through regression analysis using extensive flight simulation data and vibration sensor measurement data. Simultaneously, a pre-set severity alarm threshold is set. Only when the real-time calculated comprehensive severity score is triggered... Only when the severity threshold is exceeded will the system determine that there is a risk of unplanned aircraft jitter, thereby activating the next level of data compensation and frequency surge process. The real-time flight attitude severity quantification array is a numerical vector containing various sub-indicators and the total score, such as [normalized cumulative rate, average delay, comprehensive severity score].

[0076] In a specific embodiment of the present invention, the severity alarm threshold The typical range for the threshold value is usually set between 0.300 and 0.450, with a typical operating value of 0.400 based on a safety margin. The core physical and hardware mechanisms underlying this setting are rooted in the flutter critical inflection point of the helicopter's mounting structure and the quantum phase-locked tracking band limit of a high-precision optical magnetometer. As a dimensionless total score derived by weighting and normalizing the dynamic cumulative rate of angular acceleration and the absolute delay time of mechanical action, when its score is below 0.300, the airframe deformation is within the normal linear elastic buffer range of airflow disturbance and mechanical linkages during normal cruise, and the sensor's background filtering and attitude compensation are sufficient to maintain a stable signal-to-noise ratio; however, once the score is below 0.300, the airframe deformation is within the normal linear elastic buffer range of airflow disturbance and mechanical linkages during normal cruise, and the sensor's background filtering and attitude compensation are sufficient to maintain a stable signal-to-noise ratio; Approaching or exceeding the critical value of 0.400 clearly indicates that the helicopter's sharp yaw or large pitch maneuvers, performed to track abnormal high-gradient trajectories, have triggered unplanned high-frequency torsional flutter (elastic whiplash effect) in the fuselage boom or tail boom structure. The rate of change of the transient spatial cutting angle of the airborne probe relative to the Earth's magnetic field lines caused by this high-frequency flutter will approach the dynamic tracking bandwidth limit of the Larmor resonant frequency phase-locked loop circuit inside the optical pump magnetometer. This can easily induce instantaneous "defocusing," brief loss of lock-on, or huge eddy current background noise generated by the cutting of magnetic field lines by the magnetically conductive metal components in the optical detector signal. Therefore, by strictly controlling the dimensionless threshold within the physical critical range of 0.300 to 0.450, the system not only filters out normal vibrations during stable flight, but also provides the system with a priori predictive advantage of tens of milliseconds before destructive electromechanical coupling interference breaks through the sensor's measurement baseline and causes substantial breaks in physical data. This allows for precise triggering of underlying hardware interruptions, enabling the system to preemptively capture characteristic peak values ​​using a 10,000 Hz-level extreme saturation clock sampling. This effectively achieves safe closed-loop control that crosses the dynamic boundaries of the flight platform and the physical limits of quantum sensing.

[0077] For example, the instruction sequence issued by S4 is received. The scenario is as follows. Assume the instruction is executed at system time 15:000. The flight control and detection integrated processing unit was activated at 15,000. Up to 15.200 200 Listening was conducted within the window. The torque sensor on the tail rotor servo connection shaft was detected at 15.018. When the reading begins to deviate from the static value, the absolute action delay offset time of the tail rotor is 15.018 - 15.000 = 0.018. Within this window, the cumulative rate of dynamic change in yaw servo angular velocity is calculated by processing the tail rotor servo encoder data. The system retrieves a preset normalization constant, assuming a normalization benchmark for the delay time. It is 0.05 Normalized benchmark for accumulative rate of angular acceleration for Weighting coefficients and The scores are 0.6 and 0.4 respectively. The system substitutes the data into the aircraft's high-maneuver attitude assessment example model for calculation, and obtains the comprehensive severity score: Finally, the system integrates these quantification results into a real-time flight attitude abrupt change quantification array with specific values ​​of [0.48, 0.36, 0.432], and outputs this array to the next step for timing compensation calculation.

[0078] The clock limit latch instruction generation module uses a real-time flight attitude drastic measurement array to determine the instantaneous hardware detection signal distortion attenuation factor, generates an airborne probe limit clock frequency latch instruction, and achieves tomographic data restoration by forcibly writing back to trigger limit saturation clock sampling.

[0079] In a specific embodiment of the present invention, the distortion attenuation factor of the instantaneous hardware detection signal is determined by using a real-time flight attitude drastic quantification array, and an airborne probe limit clock frequency latch instruction is generated. This includes: translating the real-time flight attitude drastic quantification array and substituting it into a pre-stored rotor attitude cutting magnetic field line background noise mapping table for retrieval, and calculating and determining the distortion attenuation factor of the instantaneous hardware detection signal.

[0080] Based on the instantaneous hardware detection signal distortion attenuation ratio, an exponential frequency offset correction correlation multiplication factor is adjusted to generate an airborne probe limit clock frequency latching instruction.

[0081] Specifically, after generating the real-time flight attitude severity quantification array, the integrated flight control and detection processing unit immediately initiates the final closed-loop feedback adjustment step. The processing unit first translates the comprehensive severity score in the real-time flight attitude severity quantification array into a data format, using it as an index to retrieve the data in parallel from a pre-stored rotor attitude cutting magnetic field line background noise mapping table in non-volatile memory. This mapping table records the interference signal intensity generated by the high-speed rotation of the rotor cutting the geomagnetic field lines under different helicopter maneuver intensities. By looking up the table, the system can calculate and determine the instantaneous hardware detection signal distortion attenuation factor caused by the sudden maneuvering and tumbling generated by the helicopter executing the preceding follow-up flight command. Next, based on the retrieved instantaneous hardware detection signal distortion attenuation factor, the system precisely adjusts an exponential frequency offset correction correlation multiplication factor. This factor dynamically adjusts the clock frequency inside the detection hardware, its physical purpose being to significantly shorten the read / write sleep cycle of the physical sensing probe during data acquisition. Based on this factor, the system finally generates an airborne probe limit clock frequency latch command. To achieve the most direct and highest-priority control, the system bypasses the conventional data bus and uses a dedicated hardware interrupt response channel to forcibly write back the airborne probe's extreme clock frequency latch instruction to the hardware register of the original trigger timing gate array of the optical pump magnetometer. This action suspends the magnetometer's original, uniform sampling timing control in real time. This control blocking manifests in the underlying circuitry as follows: the system sends a hardware-level shielding signal to the magnetometer's FPGA's internal timer reset pin or clock enable pin via a physical direct connection, forcibly suspending the current normal trigger queue that relies on the internal crystal oscillator's average speed counting. Simultaneously, a write operation activates the latch pin storing the full-frequency configuration factor, thereby triggering the physical sensing probe to forcibly activate a far-than-normal extreme saturation clock sampling mode at a specific moment when the transient angular acceleration of the spacecraft exceeds the steady-state threshold. This allows for the precise capture of tomographic data that is easily truncated due to violent maneuvers, achieving data restoration and forced filling of the target area's spatial topological continuity.

[0082] The first formula above describes the instantaneous hardware detection signal distortion attenuation factor. The process of obtaining it. It is a dimensionless value, determined by consulting the rotor attitude cutting magnetic field line background noise mapping table. The query index is the total attitude severity score calculated in the previous steps. The second formula calculates the new sampling frequency. . This is the original sampling frequency, in units of . It is achieved by multiplying the original frequency by a factor... driving exponential term The exponential function was obtained. Its use ensures that even... Even a small increase can lead to a significant increase in frequency. This is a preset dimensionless frequency surge sensitivity coefficient used to adjust the drasticness of frequency changes, ensuring the dimensional consistency of the expression. The third formula expresses the airborne probe's limiting clock frequency latching command. The generation of this instruction is for a new sampling frequency. The result of hardware trigger logic encoding is in the form of a specific register address and write value pair.

[0083] In one specific embodiment of the present invention, It is a preset dimensionless frequency surge sensitivity coefficient used to adjust the drasticness of frequency changes. In actual high-precision airborne geophysical exploration projects, The typical value range is usually set between 1.2 and 1.8, with a typical operating value calibrated at 1.5. Its core mathematical basis and engineering boundary are deeply coupled with the nonlinear signal attenuation law caused by helicopter maneuvers and the physical throughput limit of the underlying hardware of the airborne quantum detector: due to the high-frequency flutter of the pod caused by high-G maneuvers, the polarization signal of the optical pump magnetometer will experience a very strong exponential degradation cliff in terms of defocus and signal-to-noise ratio. To offset this nonlinear high-frequency data loss, the hardware-level clock sampling frequency must be increased exponentially, either equally or even closer to the boundary. When the coefficient is calibrated to 1.5, the instantaneous distortion attenuation factor in the conventionally measured range of 1.0 to 2.0 can be represented. Through the base This is smoothly and rapidly mapped to a frequency surge multiplier of approximately 4.5 to 20 times. The setting of this key sensitivity factor ensures both the ability to generate high-density saturation sampling frequencies within tens of milliseconds of maneuvering to fill spatial physical gaps, and also rigorously prevents... Excessively large values ​​(such as >2.0) can lead to serious malfunctions such as exponential explosions triggered by minor normal vibrations, instantaneous breakdown of the probe's maximum write frequency causing data bus overflows, or thermal failure of components. Thus, a precise nonlinear mathematical barrier is established between maximizing hardware performance and maintaining the physical safety envelope of the system.

[0084] The rotor attitude cutting magnetic field line background noise mapping table is a static lookup table derived from extensive flight experiments and electromagnetic simulations. Its index is the attitude severity score, and the value is the corresponding signal attenuation factor. The frequency offset correction correlation multiplication factor is a core control parameter used to calculate the new sampling period; its exponential form is designed to achieve a non-linear, responsive frequency boost. The airborne probe limit clock frequency latch instruction is an override control command designed to directly overwrite the sensor's underlying timer and trigger hardware settings. The dedicated peripheral interrupt response channel is a physically independent hardware connection line with the highest processing priority. It bypasses the software protocol stack that requires queuing and arbitration, ensuring immediate instruction effectiveness. The ultimate goal of this step is to achieve non-uniform detection: using a conventional frequency to save energy and storage when the flight attitude is stable, and instantaneously boosting the sampling frequency to the hardware limit during moments of drastic flight attitude changes and when data is most prone to distortion, thereby forcibly filling in any missing data.

[0085] In a specific embodiment of the present invention, the physical mechanism of the background noise generated by the rotor attitude cutting magnetic field lines is as follows: the transmission components and some blade structures of the helicopter rotor often contain magnetically conductive metal materials. When the helicopter performs a large maneuver, causing the rotor disk plane to tilt violently, the high-frequency flapping of the blades and the periodic pitch-changing action will cause changes in the local induced eddy currents around the rotor. This, in turn, will periodically interfere with or dynamically shield the geostatic magnetic field lines received by the detector under the fuselage. This high-frequency electromagnetic environment oscillation will be directly reflected as distortion of the detection signal and a surge in background noise. Through this mapping table, the intensity of the mechanical maneuver can be quantitatively converted into the attenuation factor caused by electromagnetic interference.

[0086] For example, the system receives a real-time flight attitude severity quantification array output by S5, where the comprehensive severity score is... The system uses this as an index to query the rotor attitude cutting magnetic field line background noise mapping table, and finds the corresponding instantaneous hardware detection signal distortion attenuation factor. Assume the system has a preset frequency surge sensitivity coefficient. The original sampling frequency of the optical pump magnetometer is 1.5. For 500 The system adjusts the frequency offset correction correlation multiplication factor accordingly and calculates the new target sampling frequency: = The system rounds the frequency to 5512. This is then encoded into a specific airborne probe limit clock frequency latch instruction. For example, this instruction might be a pair of data: [register address: Write the value: Finally, the system forces the instruction back to the magnetometer's timing gate array via the SPI or I2C hardware interrupt interface, forcing its original 500... The sampling timing was overwritten; immediately switch to approximately 5.5. High-energy sampling is performed to achieve dynamic compensation and precise capture of geophysical data during intense maneuvers.

[0087] In a specific embodiment of the present invention, the tomographic data is restored by forcibly writing back the trigger limit saturation clock sampling, including: forcibly writing back the airborne probe limit clock frequency latch instruction in reverse through the peripheral dedicated interrupt response channel, thereby blocking the original trigger timing gate array of the optical pump magnetometer.

[0088] This forces the physical sensing probe to activate extreme saturation clock sampling the instant the transient angular acceleration of the space body exceeds the steady-state threshold, capturing truncated tomographic data to restore the spatial topological coherence of the target area.

[0089] Specifically, after guiding the helicopter to the preset aerial survey starting point, the integrated flight control and detection processing unit first sends wake-up and configuration commands to the onboard sensor array via the underlying communication bus, activating the onboard induction coil and optical pump magnetometer equipment. The controlled equipment then switches to and enters a continuous sampling mode uninterrupted by conventional waypoints, continuously scanning the subsurface medium to acquire the transient signal stream of the underlying physical field within the target area in real time. This signal stream is essentially a set of discrete digital signals in a one-dimensional time series, containing not only weak high-frequency signals generated by the geological bodies in the target area but also strongly superimposed the extremely strong, slowly varying static magnetic field background of the Earth and a broad regional geological tectonic field. To remove these macroscopic background interferences that obscure real geological anomalies, the system imports the received transient signal stream of the underlying physical field into the hardware processing engine in the form of lossless data frames. The engine has a dedicated digital signal processing pipeline logic pre-installed, primarily using high-pass filtering with a preset cutoff frequency or directly performing moving average background removal operations to perform real-time dynamic removal processing on the input continuous signal stream using a sliding window approach.

[0090] In one specific embodiment of the present invention, to address the risk of servo control disconnection caused by severe phase shift and group delay due to excessively long data windows in traditional high-frequency extraction algorithms using high-order FIR filters, the present invention preferably employs a low-delay infinite-long unit impulse response (IIR) filter algorithm with phase lead compensation, or an adaptive background elimination algorithm based on the Kalman predictive state equation. While ensuring a forced pull-down "zeroing" of the 10,000-tesla-level static DC substrate, the algorithm controls the group delay within the range set by the current detection system. Within a time step of 5ms to 20ms, the physical feasibility of "zero-delay" closed-loop control from detection to servo execution is truly guaranteed.

[0091] This processing procedure suppresses extremely low-frequency DC references and inhibits the slowly changing static Earth's main magnetic field and regional field background signals from the transient physical field signal stream sampled from the original hardware layer. After background stripping, the system successfully discards the spatially spreading redundant data field, retaining only the high-frequency disturbances caused by underground geological anomalies during rapid helicopter fly-throughs. This allows for the precise extraction of transient physical field characteristic parameters, which are then used as the original, pure data for subsequent assembly of the three-dimensional spatial gradient matrix, unaffected by regional baseline drift.

[0092] In the above formula, The real-time intensity of transient signal streams of the underlying physical field captured by airborne detection equipment, measured in nanoteslas. ; This represents the static Earth's main magnetic field and regional field background signal superimposed on the signal and slowly changing with space; This indicates a weak but rapidly changing effective local anomaly signal caused by geological anomalies, while This represents the Gaussian white noise generated by the instrument itself and the environment. The second formula details the method of extracting high-frequency disturbances using a moving average background removal operation. This refers to the time rate of change component of the estimated transient characteristic parameter of the physical field, in units of... ; This refers to the time interval between adjacent sampling points in continuous sampling mode, in units of... The third formula illustrates a high-order method for anomaly extraction based on finite-length unit impulse response (FIR) high-pass filtering. Represents the transient characteristic parameters of the physical field obtained after time-domain digital filtering, with units of 1200 ppm. ; These are the design tap coefficients for a high-pass filter, dimensionless, used to cut off low-frequency trends; This refers to the filter order. Both of these mathematical operations forcibly cancel or block low-frequency, high-amplitude values ​​in the time domain. The composition ensures that the signal-to-noise ratio is optimized at the physical dimension scale during subsequent matrix solving.

[0093] In one specific embodiment of the present invention, These are the dimensionless tap coefficients for designing a high-pass filter to cut off low-frequency trends. In practical high-precision magnetic measurement engineering for airborne applications, this is used for sampling rates as high as kilohertz and with high order... High-order linear-phase FIR filters, typically set between 512 and 2048, have center tap coefficients... Typical values ​​for this are usually very close to positive 1 (e.g., 0.998 to 0.999), while the remaining tap coefficients are symmetrically distributed on both sides. The envelope decay exhibits a small negative value in the Sinc function range from -0.1 to 0, and must strictly satisfy the condition that the sum of all tap coefficients is always equal to 0 (i.e., ...). The mathematical conditions of the coefficient quantification are based on the deep cross-border integration of the helicopter cruise ground speed kinematic model and the heterogeneous spatial wavelength scale of the geophysical field: the blind veins or faults tracked by aerial exploration often belong to local geological bodies with a spatial span of tens to hundreds of meters. When the helicopter travels at approximately When the economy speeds by and cuts through, it will generate approximately [a certain amount of energy] in the time domain. Up to 10 The high-frequency abrupt change in the physical field of the frequency band; while the geomagnetic main magnetic field with an intensity of over 50,000 nanoteslas and the deep regional tectonic basement exhibit extremely slow variation characteristics over tens of kilometers in space, which are reflected in the time domain as less than 0.05. The ultra-low frequency "DC drift" background. Therefore, based on the extremely low low-frequency cutoff frequency (e.g., set to 0.05), This complete set of methods, rigorously calculated and solidified using the window function method, is The coefficient array can construct a digital filter "funnel" with an extremely steep transition band. Through the discrete time-domain convolution physical operation of "full retention of the center point and wide-area negative phase cancellation of the long hysteresis window", it can instantly pull down and "clear" the static DC substrate of 10,000 nanotesla level. Thus, without causing obvious phase delay of high-frequency signals and spatial topological distortion, it can effectively peel off and extract the micro-nanotesla level Gaussian transient anomalies superimposed on it, which are extremely easy to be submerged.

[0094] Among them, the airborne induction coil and optical pumped magnetometer are the core hardware combination in high-precision airborne geophysical exploration. The induction coil is extremely sensitive to the instantaneous rate of change of the magnetic field gradient, while the optical pumped magnetometer can provide extremely high-resolution measurements of the total absolute magnetic field or its three components based on the principle of quantum energy level transitions. The continuous sampling mode differs from traditional spatial fixed-distance, fixed-point triggered sampling. It requires the detection equipment to maintain continuous data generation at a fixed and extremely high frequency (e.g., thousands of hertz) during flight. The purpose is to prevent the loss of sharp-scale geological anomaly targets due to sampling gaps during flight maneuvers. The transient signal stream of the underlying physical field contains the total superposition effect of all magnetic field sources within a specific depth below the helicopter extending upwards to the flight altitude. High-pass filtering or moving average background removal is a classic real-time method for separating the geophysical background field. Its physical basis is that the frequency of sensor signal changes caused by the helicopter flying at high speed over small-scale, high-fault gradient anomalies is significantly higher than the time-varying frequency caused by the large-scale smooth superposition of the Earth's own background field. The transient characteristic parameter of the physical field refers to the residual high-energy anomaly signal after removing the macroscopic spatial geological background and the inherent drift interference of the Earth's main magnetic pole. It eliminates the masking of the absolute reference value and directly maps the physical property heterogeneity boundary between the local shallow crust or hidden rock mass and the surrounding rock directly below the helicopter.

[0095] For example, after the helicopter enters the target area boundary, the integrated flight control and detection processing unit immediately issues a hardware takeover command, in which the optical pumped magnetometer is activated and forced into continuous sampling mode. At this time, the system sets its sampling frequency to 1000. That is, the time interval between two adjacent samplings. for At a certain sampling moment during the cruise process The current frame data value of the transient signal stream of the underlying physical field read by the sensor through the high-speed ADC is... Due to the influence of regional geomagnetic latitude and locally gentle geological formations, the background static magnetic field of the Earth's main magnetic field in this airspace remains approximately at [value missing]. Nearby, and exhibiting a slow drift on the order of several seconds. At the immediately preceding fraction of a second sampling time... The sensor digital reading is To avoid cumbersome nonlinear fitting of the regional field, the engine's internal FPGA hardware solution module instantly executes the basic background subtraction formula to obtain the pure outlier values ​​of the physical field after stripping away the macroscopic background: The system substitutes the numerical values ​​and calculates the instantaneous anomaly: Through this low-resource-consumption, microsecond-level, and extremely low-latency background stripping calculation, the regional background DC, originally as high as 54,000 nanoteslas, was instantly subtracted, filtered, and cleared to zero. The system accurately captured and preserved the instantaneous high-frequency signal response amplitude excited upward by potentially high-magnetic-susceptibility concealed ore bodies underground. This amplitude ranges from a few tenths to several hundred nanoteslas. The pure mutation field strength of the order of magnitude was successfully extracted and officially used as a core three-dimensional basis component of the transient characteristic parameter of the physical field generated in this cycle. Together with the parameters extracted by the other array probes, it was sent to the next level logic gate sequence to solve the dimensionally tight spatial gradient tensor pulse.

[0096] The above content is merely an example and illustration of the concept of the present invention. Those skilled in the art can make various modifications or additions to the specific embodiments described, or use similar methods to replace them, as long as they do not deviate from the concept of the invention or exceed the scope defined by the present invention, and all such modifications and additions should fall within the protection scope of the present invention.

Claims

1. A helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance, characterized in that, include: The frontal spatial gradient tensor pulse generation module acquires the physical field transient signal stream sampled by the original hardware layer collected by the helicopter airborne probe, and performs transient three-dimensional differential calculation and singular value decomposition through the hardware calculation engine to generate frontal spatial gradient tensor pulses. The geophysical morphological feature parameter generation module extracts the physical anomaly gradient magnitude vector and the main feature evolution direction from the frontal spatial gradient tensor pulse, and combines them to generate geophysical morphological feature parameters. The equation generation module obtains the helicopter's current flight attitude package and extracts flight dynamics extreme boundary data. It then performs forced coverage and assimilation by combining geological and physical morphological parameters to generate a set of aerodynamic control equations. The flight operation command sequence generation module inputs the update time constant of the current space physics exploration node into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, generate a geological morphology space-following flight operation command sequence and issue it; The flight attitude severity quantification array generation module monitors the real-time mechanical response force sequence during the execution of the geological morphology space follow-up flight operation command sequence, and generates a real-time flight attitude severity quantification array by combining the aircraft composite maneuver attitude evaluation example model. The clock limit latch instruction generation module uses a real-time flight attitude drastic measurement array to determine the instantaneous hardware detection signal distortion attenuation factor, generates an airborne probe limit clock frequency latch instruction, and achieves tomographic data restoration by forcibly writing back to trigger limit saturation clock sampling.

2. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The process of acquiring the transient physical field signal stream sampled by the helicopter's airborne probe through the hardware layer, and performing transient three-dimensional differential and singular value decomposition through a hardware solution engine to generate a leading-edge spatial gradient tensor pulse, includes: The transient physical field signal stream sampled from the original hardware layer is imported into the hardware solution engine to perform transient three-dimensional differential calculation and extract transient feature parameters of the physical field. Calculate the partial derivatives of the transient characteristic parameters of the physical field in the orthogonal triaxial directions and assemble them into a three-dimensional spatial gradient matrix; Singular value decomposition is performed on the three-dimensional spatial gradient matrix to obtain the leading spatial gradient tensor impulse.

3. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The extraction of the physical anomaly gradient magnitude vector and the evolution direction of the main features from the leading edge spatial gradient tensor pulses are combined to generate geophysical morphological feature parameters, including: The sum of squares of all diagonal elements of the singular value diagonal matrix in the leading spatial gradient tensor impulse is obtained and the square root operation is performed to obtain the total absolute variability as the physical anomaly gradient magnitude vector. The feature space analytical comparison of the leading spatial gradient tensor impulse is performed, and a set of orthogonal feature vectors corresponding to the maximum eigenvalues ​​are extracted as the main feature evolution direction. By combining the gradient magnitude vector of physical anomalies with the evolution direction of the main features, geological and physical morphological feature parameters are generated.

4. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The process involves acquiring the helicopter's current flight attitude package and extracting flight dynamics extreme boundary data, then combining this data with geophysical morphological parameters for forced coverage and assimilation to generate a set of fundamental aerodynamic control equations, including: Read the helicopter's current flight attitude packet and extract the rotor's maximum tilt angle limit and the tail rotor's minimum available yaw thrust from the physical extreme value envelope table built into the helicopter's flight control system, and combine them into flight dynamics extreme value boundary data; The physical anomaly gradient magnitude vector attached to the geophysical morphological feature parameters is extracted as the reciprocal penalty factor to forcibly cover the basic forward velocity scalar mapping formula. The evolution direction of the main features in the geophysical morphological parameters is forcibly assimilated into the main guiding input source of the tail rotor aerodynamic yaw rate. Combined with the basic forward velocity scalar mapping formula that has been covered and changed and the extreme boundary data of flight dynamics, the aerodynamic control basic solution equation set is obtained by integration and recombination.

5. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The process of inputting the current space physics exploration node update time constant into the aerodynamic control fundamental equation set for online numerical iterative convergence solution, generating and issuing a sequence of geological morphology space-following flight operation commands, includes: The current space physics exploration node update time constant is input into the aerodynamic control fundamental solution equation set to perform online numerical iterative convergence solution, and the periodic pitch compensation amount and tail rotor yaw angle compensation degree are obtained. Based on the periodic pitch compensation, a forced anti-run-out deceleration command is generated; based on the tail rotor yaw angle compensation, a tail rotor heading follow-up yaw operation intervention command is generated. The forced anti-runaway deceleration command and the tail rotor heading follow-up deflection operation intervention command are encoded with overlapping bit mask and merged into a sequence of geological morphology space follow-up flight operation commands for issuance.

6. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The real-time mechanical response force sequence during the monitoring and execution of the geological morphology spatial follow-up flight operation command sequence, combined with the aircraft composite maneuver attitude evaluation example model, generates a real-time flight attitude severity quantification array, including: The system continuously monitors the tension and friction feedback waveforms returned by sensors on the main rotor periodic pitch control rod and the tail rotor drive servo connection shaft of the helicopter to obtain the real-time mechanical response force sequence within a set timing rolling window. The cumulative rate of dynamic change in yaw servo angular velocity and the absolute motion delay offset time of each mechanical operating surface are calculated by summing the forces contained in the real-time mechanical response force sequence. By substituting the cumulative rate of dynamic change in yaw servo angular velocity and the absolute action delay offset time into the aircraft's high-maneuver attitude evaluation case model, a real-time flight attitude severity quantification array is generated.

7. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The step of using a real-time flight attitude drastic quantification array to determine the instantaneous hardware detection signal distortion attenuation factor and generating an airborne probe limit clock frequency latch instruction includes: The real-time flight attitude drastic quantification array is translated and substituted into a pre-stored rotor attitude cutting magnetic field line background noise mapping table for retrieval, and the instantaneous hardware detection signal distortion attenuation factor is calculated and identified. Based on the instantaneous hardware detection signal distortion attenuation ratio, an exponential frequency offset correction correlation multiplication factor is adjusted to generate an airborne probe limit clock frequency latching instruction.

8. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 1, characterized in that: The method of restoring tomographic data by forcibly writing back to trigger extreme saturation clock sampling includes: The onboard probe's limit clock frequency latch instruction is forcibly written back in reverse through the dedicated interrupt response channel of the peripheral device, thus blocking the original trigger timing gate array of the optical pump magnetometer. This forces the physical sensing probe to activate extreme saturation clock sampling the instant the transient angular acceleration of the space body exceeds the steady-state threshold, capturing truncated tomographic data to restore the spatial topological coherence of the target area.

9. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 2, characterized in that: The transient physical field signal stream sampled from the original hardware layer is imported into the hardware solution engine to perform transient three-dimensional differential calculation, extracting transient feature parameters of the physical field, including: The airborne multi-probe induction coil and optical pump magnetometer, arranged in a three-dimensional spatial orthogonal array, are activated to enter continuous sampling mode and acquire the physical field transient signal flow of the original hardware layer sampling of the continuous waveform within the basic target area. By using high-pass filtering or moving average background removal operations, the slowly changing static Earth's main magnetic field and regional field background signals are effectively filtered out from the physical field transient signal stream sampled from the original hardware layer. The high-frequency disturbances caused by underground geological anomalies are retained, and the transient characteristic parameters of the physical field are extracted.

10. The helicopter retrospective aerial survey planning system based on geophysical anomaly target area guidance according to claim 5, characterized in that: The step of inputting the current space physics exploration node update time constant into the aerodynamic control fundamental equation set to perform online numerical iterative convergence solution, and obtaining the periodic pitch compensation and tail rotor yaw angle compensation, includes: Based on the difference between the target forward velocity and the current actual velocity calculated from the aerodynamic control fundamental equation set, the periodic pitch compensation amount is calculated by referring to the preset helicopter inverse dynamics model data table. The required angular acceleration is calculated by using the difference between the target yaw rate and the current actual yaw rate obtained from the aerodynamic control fundamental equations. Combined with the pre-set helicopter tail rotor aerodynamic efficiency inversion function mapping table, the tail rotor yaw angle compensation degree is obtained.

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