A ground laboratory remote sensing flight simulation platform and control method

Abstract

The present invention relates to a ground laboratory remote sensing flight simulation platform and a control method. A ground laboratory remote sensing flight simulation platform, wherein a seat rail, a floor, a wiring box, a non-airtight opening airtight cover, a downward optical window, a camera cabin door and a non-airtight protective door manual opening mechanism, and a satellite navigation signal simulator are installed inside the cabin; a truss is connected to the bottom of the fuselage, a pod is connected to the truss, and an antenna cover is installed on the outer surface of the fuselage; a fuselage mounting bracket is installed below the fuselage structure, and a boarding ladder is installed at the side door of the fuselage; one end of the power system is installed outside the fuselage, and the other end is installed inside the cabin; the present invention can complete the installation, airworthiness test, joint debugging test, mission planning, simulated flight and other tasks of various remote sensing equipment in the laboratory by building a remote sensing flight simulation platform.

Description

Technical Field

[0001] The invention relates to a flight simulation platform, in particular to a ground laboratory remote sensing flight simulation platform and a control method. Background Art

[0002] One of the preliminary tasks of an aerial remote sensing mission is to install the remote sensing equipment on the aircraft, and then conduct an onboard ground power-on test on the remote sensing equipment to ensure that the equipment works properly during the actual flight. The usual onboard ground power-on test first requires technicians to travel to the airport where the remote sensing aircraft is located, go through a series of complex approval procedures to enter the airport, and then install the remote sensing equipment on the aircraft before conducting a power-on test. When technicians are performing installation, power-on, and debugging operations on the aircraft, they will inevitably encounter unexpected technical problems. Since they are out of town or too far away from the laboratory, they cannot get technical support quickly, so the problem cannot be solved quickly, resulting in delays in the mission schedule and a waste of a lot of resources. Summary of the invention

[0003] The purpose of the present invention is to provide a ground laboratory remote sensing flight simulation platform and control method, so that technicians can complete the on-board ground power-on test in the laboratory, solve the problems arising in the work process as quickly as possible, simulate the environment during actual flight, and ensure the smooth completion of remote sensing tasks during actual flight, thereby improving work efficiency, reducing losses, and saving resources.

[0004] In order to achieve the above object, the present invention has the following technical solutions:

[0005] A ground laboratory remote sensing flight simulation platform of the present invention comprises:

[0006] Fuselage, seat rail, floor, wiring box, non-airtight opening, downward optical window, optical glass, manual opening mechanism of camera door, manual opening mechanism of non-airtight protection door, satellite navigation signal simulator, power divider, pod, truss, radome, fuselage mounting bracket, boarding ladder, power system, platform center of gravity position; among which, seat rail, floor, wiring box, airtight cover of non-airtight opening, downward optical window, manual opening mechanism of camera door and non-airtight protection door, and satellite navigation signal simulator are installed inside the cabin; truss is connected to the bottom of fuselage, pod is connected to truss, and radome is installed on the outer surface of fuselage; fuselage mounting bracket is installed under fuselage structure, and boarding ladder is installed at the side door of fuselage; one end of power system is installed outside fuselage, and the other end is installed inside cabin;

[0007] The fuselage starts from the rear end of the cockpit and ends at the front end of the tail cabin, including the structure below the lower edge of the aircraft porthole, including the structure with the floor facing downward and the structure with the floor facing upward to the lower edge of the porthole;

[0008] The seat floor rail is installed inside the cabin and fixedly connected to the floor, and is used to install seats, cabinets, control consoles, and remote sensing equipment;

[0009] The floor is laid on the aircraft floor beam and is fixed to the floor beam by screws;

[0010] The wiring box is installed on the left and right sides and the middle position of the cabin. The wiring boxes on the left and right sides are fixedly connected to the side walls of the cabin, and the wiring box in the middle is fixedly connected to the floor. It is used for laying cables between mission equipment. The wiring box can prevent the cables from being stepped on and at the same time take into account the aesthetic requirements; openings are reserved on the inner sides of the wiring boxes on the left and right sides as channels for the mission equipment cables to enter the wiring boxes. The openings can be blocked with plugging covers, and a top cover is reserved on the top of the wiring box to facilitate opening and laying cable components;

[0011] The non-airtight opening is located in the tail cabin of the aircraft. Inside the airtight cover, a mounting platform parallel to the horizontal reference plane of the aircraft is arranged above the non-airtight opening. The mounting platform has mounting holes, and the remote sensing equipment and the airtight cover can be installed at the same time. The airtight cover plays an airtight role and is installed on the mounting platform to isolate the aircraft airtight cabin from the environment outside the cabin. The non-airtight opening is located at the outer skin of the belly of the aircraft, and the opening size is 700×700 mm measured in the flight direction×wing span direction.

[0012] There are three downward-looking optical windows located at the bottom of the cabin, and a stable platform for installing a camera is provided. The stable platform is a rectangular parallelepiped structure, and chamfers are provided at the four corners of the rectangular parallelepiped structure. A cover plate is provided on the top of the stable platform, and the cover plate is fixedly connected to the rectangular parallelepiped structure. A handle is provided on the cover plate, which can be used during installation and removal; a through hole is provided in the central part of the rectangular parallelepiped structure for the downward-looking optical channel of the camera;

[0013] The camera is mounted on a stable platform when in use, and the stable platform is mounted on a structural mounting platform;

[0014] The maximum weight that the structural mounting platform can bear is less than 150kg. The structural mounting platform is provided with four threaded holes for mounting corresponding to the four mounting holes of the stable platform. The geometric center axis of the optical glass in the downward optical window coincides with the hole axis of the stable platform. The distance between the mounting surface of the structural mounting platform and the upper surface of the optical glass should be less than 150mm. The parallelism between the mounting surface of the structural mounting platform and the upper surface of the optical glass is not greater than 1 degree. The mounting surface of the structural mounting platform is parallel to the horizontal plane of the aircraft.

[0015] There are three optical glasses, all of which are located on the three optical windows below the structural mounting platform. The effective aperture of the three optical glasses is ≥535mm×525mm measured in the front-to-back direction×the wingspan direction;

[0016] The manual opening mechanism of the camera cabin door is located below the cabin floor and is fixedly connected to the inner wall of the cabin bottom; the manual opening mechanism of the non-airtight protection door is located in the tail cabin of the aircraft and is fixedly connected to the tail cabin floor;

[0017] The camera door is located below the optical glass and is used to protect the camera lens and optical glass installed on the downward optical window, and to protect the camera lens and optical glass from being hit by gravel, ice, rain, snow or something brought up from the runway by the landing gear during takeoff and landing;

[0018] The non-airtight protection door is located below the non-airtight opening and close to the skin, and is used to protect the remote sensing equipment from damage during takeoff and landing of the aircraft; the non-airtight protection door is closed during takeoff and landing, and opened when the remote sensing equipment is operating; the non-airtight protection door can ensure the observation field angle of the mission equipment;

[0019] The satellite navigation signal simulator is fixedly connected to the side wall of the aircraft cabin; the power divider is fixedly connected to the side wall of the aircraft cabin; the power divider is used to provide satellite navigation signals for various remote sensing devices;

[0020] There are four trusses, two of which are installed on the left and right sides of the aircraft fuselage, with their upper ends fixedly connected to the side of the aircraft and their lower ends fixedly connected to the pod; the installation interface between the aircraft fuselage structure and the truss structure is connected by bolts;

[0021] The pods are installed on both sides of the fuselage of the aircraft through trusses, one on each side, with a center distance of 4.2 meters and an error of no more than 0.1 meters. The relative deformation of the two antennas installed in the pods is required to be less than 0.1 mm during stable flight;

[0022] There are three radomes in total, one radome is arranged under the belly of the fuselage, and a left radome and a right radome are arranged on both sides of the fuselage;

[0023] The function of the camera and lighting device: When the radome is closed, the movement of the antenna can be observed at any time, and a location can be selected for installation according to the aircraft conditions;

[0024] The fuselage mounting bracket is located below the fuselage structure and connected to the fuselage belly;

[0025] The boarding ladder is located at the boarding gate on the left side of the fuselage and is used for test personnel to board the aircraft. The load capacity is 300Kg;

[0026] The power supply system comprises a laboratory wall power supply, a power conversion device, a fuselage external grounding interface, a power distribution cabinet, and a power distribution box. The power conversion device is located near the laboratory wall power supply, the fuselage external grounding interface is installed on the fuselage structure surface, the power distribution cabinet is located at the rear end of the cabin cockpit, the power distribution box comprises a power distribution box 1 (P1), a power distribution box 2 (P2), and a power distribution box 3 (P3), the power distribution box 1 (P1), the power distribution box 2 (P2), and the power distribution box 3 (P3) are all located inside the cabin, the power distribution box 1 (P1) is located in the front equipment area, the power distribution box 2 (P2) is located in the rear equipment area, and the power distribution box 3 (P3) is located in the tail cabin equipment area, and the three distribution boxes are all fixedly connected to the side wall of the cabin;

[0027] The center of gravity position of the platform changes with the different installation positions of the equipment on the platform, and the center of gravity position of the platform in different installation states needs to be calculated; obtaining an accurate center of gravity position of the platform can ensure the smooth flight of the aircraft, thereby ensuring the quality of remote sensing data acquisition.

[0028] The location layout of the remote sensing equipment determines the center of gravity of the platform; after the remote sensing equipment acquires GNSS data, it is necessary to calculate the GNSS eccentricity component based on its own location and the location of the satellite navigation signal simulator for post-processing of the GNSS data.

[0029] Among them, the truss includes a frame, four truss connecting rods, two cross bars, reinforcement ribs, connector one, connector two, connector three, and connector four. The four truss connecting rods are hinged at the four corners of the frame. The four truss connecting rods are used to connect the pod. The two cross bars are fixed at the two ends of the frame. The two cross bars are components for lifting. The reinforcement ribs are fixed in the frame to increase the mechanical strength of the frame. Connector one, connector two, connector three, and connector four are used to connect the fuselage.

[0030] Among them, the connecting member 1 is provided with a U-shaped frame, a T-shaped plug plate, and a bottom plate with a socket, wherein the two ends of the U-shaped frame are connected to the frame, the bottom plate with the socket is connected to the frame, and the bottom plate with the socket is located in the U-shaped frame, the T-shaped plug plate can be inserted into the bottom plate with the socket, and can be connected to the bottom plate with the socket by bolts; a screw hole is provided on the top of the T-shaped plug plate, which can be used to connect to the fuselage; the connecting member 2 is provided with a branch and a vertical plate, the branch is connected to the middle of the frame, the bottom of the vertical plate is connected to the branch, and the vertical plate is provided with screw holes; the connecting member 3 is a horizontal plate, which is connected to the frame and has screw holes; the connecting member 4 is a transverse rod, and the transverse rod is provided with screw holes for connecting to the fuselage.

[0031] The pod includes a front fairing, a rear fairing and a middle section cabin, which are an integral structure. The middle section cabin includes an outer shell, an antenna cover and a connecting end plate, which are fixed on the outer shell, and both ends of the antenna cover are connected to the connecting end plate.

[0032] The functions of the camera and lighting device are: when the antenna cover is closed, the movement of the antenna can be observed at any time, and a location can be selected for installation according to the aircraft situation.

[0033] Among them, 6 mounting plates are arranged in the belly antenna cover, which are symmetrically distributed on both sides of the symmetry axis of the aircraft; the center distance between the left and right mounting plates is 600-800mm.

[0034] A control method for a ground laboratory remote sensing flight simulation platform of the present invention, wherein the satellite navigation signal simulator provides GNSS data for remote sensing equipment on the ground laboratory remote sensing flight simulation platform by simulating RF signals of multiple satellite navigation systems, and the steps are:

[0035] 1) DSP initialization, including reading simulation parameters, determining simulation time, calculating satellite navigation signal simulator trajectory, and reading satellite ephemeris parameters;

[0036] 2) Navigation message calculation: To simulate the navigation message, the broadcast ephemeris file needs to be read. This file can be downloaded from the CDDIS website of the data center under IGS. The message parameters are extracted from the broadcast ephemeris file and converted into non-polar binary code or polar binary code. Then, the parity check code is generated by the check code encoder. After being converted into a relative code, it is modulo-2 added with the folding code. Finally, a time mark is added at the end of each string to complete the encoding of the navigation message.

[0037] 3) Satellite visibility judgment;

[0038] There is a relative motion between the satellite and the earth. After a period of time, there will inevitably be a state change of rising and falling stars at the same earth coordinate point. When the satellite's pitch angle is greater than 5 degrees, the satellite is considered to be visible; when the pitch angle is less than 5 degrees, the satellite is considered to be invisible. The broadcast ephemeris file generally stores the ephemeris parameters of nearly 30 navigation satellites in different time periods. After receiving the ephemeris parameters of all satellites, the DSP calculates the position of the satellite according to the set simulation start time, and then calculates the relative position of the two according to the initial position of the satellite navigation signal simulator, thereby obtaining the pitch angle and azimuth.

[0039] In the GPS system, the measurement coordinate system is established with the phase center of the GPS receiving antenna, and the elevation and azimuth of the satellite are calculated in this coordinate system. It is known that the position of the GPS satellite in the WGS84 coordinate system at time t is X tWGS-84 =(x T ,y T , z T ) T , the target position is X = (x, y, z) T; The position of the target in the earth coordinate system can be calculated as (L, B, H) and the position of the satellite in the measurement coordinate system is The T in the footer refers to time T, and C refers to the satellite; the calculation formula for the satellite's elevation angle E and azimuth angle A is:

[0040]

[0041] in,

[0042] In the GLONASS system, the satellite position is also converted from the geocentric coordinates to the station-centric coordinate system centered on the satellite navigation signal simulator position P;

[0043] Assume [X p Y p Z p ] T is the satellite station center coordinate, [X sv Y sv Z sv ] T and [X rec Y rec Z rec ] T are the coordinates of the satellite and the satellite navigation signal simulator in the PZ-90 geocentric coordinate system, B rec and L rec are the latitude and longitude of the satellite navigation signal simulator in the geodetic coordinate system, respectively.

[0044]

[0045] In the formula, H is

[0046]

[0047] After calculating the satellite's station center coordinates, the elevation angle E and azimuth angle A of the satellite navigation signal simulator observing the satellite can be obtained:

[0048]

[0049] If E ≥ E min , E min If it is the minimum elevation angle of the satellite, the satellite is a visible satellite. In subsequent processing, it is necessary to calculate the DOP value, error parameters, and signal status parameters for this satellite. Otherwise, it is an invisible satellite.

[0050] 4) Satellite position and speed calculation;

[0051] The equation describing the satellite motion trajectory can be obtained from the ICD file (space signal interface control file) of the satellite navigation system:

[0052]

[0053]

[0054] Where (x, y, z) represents the satellite coordinates, which is a function of time t; Vx(t), Vy(t), and Vz(t) are the first-order derivatives of the satellite coordinates with respect to time, representing the satellite's velocity; It is the second-order derivative of the satellite coordinates with respect to time, indicating the acceleration of the satellite; the Earth's gravitational constant μ = 398600440m 3 / s 2 ; Reference ellipsoid major radius R e =6378136m; Gravity potential second band harmonic coefficient The Earth's rotation speed ω=7.292115×10 -5 rad / s; It is the sum of the accelerations of the sun and the moon;

[0055] The fourth-order Runge-Kutta method is used to calculate the fourth slope value k under the premise of selecting the step size. 1 , k 2 , k 3 , k 4 , and use its weighted average as the average of the weighted slope, so as to obtain the approximate values ​​of the satellite position and velocity in the interval:

[0056]

[0057] Among them, k 1 is the slope at the beginning of the time period, k 2 is the slope at the midpoint of the time period, k 3 It is also the slope at the midpoint of the time period, k 4 is the slope at the end of the time period; y(i) is a vector function; x(i) is a vector; h is the integration step;

[0058] The number of integration times n required in the entire integration process is n=|t i -t b | / h, the integration time interval under normal circumstances |t i -t b | is not an integer multiple of the step size h, so the integration step size of the last step is smaller than the selected step size, which is h l =|t i -t b |-n*h, then after the entire integration process is completed, the integration time is the required solution time t i ;

[0059] Satellite signals are affected by the rotation of the earth during propagation. The satellite position and speed calculated by the above formula are the values ​​at the time of signal reception. In order to obtain the values ​​at the time of signal transmission, the result values ​​need to be corrected:

[0060]

[0061]

[0062] Among them, X(t), Y(t), Z(t) are satellite positions; U(t), V(t), W(t) are satellite velocities; Ω e is the average angular velocity of the Earth’s rotation; t is the end time of the integration interval; t e is the front moment of the integration interval; Ω e *(tt e ) is the angle that the earth rotates during the propagation of satellite signals;

[0063] 5) Calculation of ionospheric delay error: When GNSS signals pass through the ionosphere, electromagnetic wave refraction errors will occur, thereby reducing the accuracy and precision of satellite navigation positioning. According to the Klobuchar model, the ionospheric delay error can be corrected:

[0064]

[0065]

[0066] In the formula, α′ i =α i / 0.6, β′ i =β i (i=0, 1, 2, 3); parameter α 0 , α 1 , α 2 , α 3 , β 0 , β 1 , β 2 , β 3 is the ionospheric parameter in the ephemeris file; F is the tilt factor; φ m is the electromagnetic latitude of the intersection of the satellite signal propagation path and the central ionosphere; T p is the phase value of the cosine wave; P is the period of the cosine wave; t is the local time; E is the elevation angle of the satellite relative to the satellite navigation signal simulator;

[0067] 6) Calculation of tropospheric delay error: When the GNSS signal passes through the troposphere, the carrier, spread spectrum code and navigation message produce delay errors. The Hopfield model can be used to correct the delay error:

[0068]

[0069] In the formula, K d is the wet component delay in the zenith direction; K w is the dry component delay in the zenith direction; E is the elevation angle of the satellite relative to the satellite navigation signal simulator; η d , η w are the correction coefficients for the dry and wet components of the troposphere, respectively; T k and e are the absolute temperature and water vapor pressure at standard sea level respectively; h d is the upper limit of the tropospheric dry component based on sea level; h u is the simulator height; p 0 is the atmospheric pressure.

[0070] 7) Doppler shift calculation: There is a Doppler shift between the navigation satellite and the satellite navigation signal simulator. In order to restore the real satellite signal, the satellite navigation signal simulator needs to calculate the Doppler shift. According to the existing formula:

[0071]

[0072] Where v is the satellite speed, λ is the signal wavelength corresponding to the signal transmission frequency f, c is the speed of light, and β is the signal incident angle, that is, the angle between the receiving end and the signal incident direction; assuming that the signal frequency transmitted by the satellite is f c , the radial velocity in the direction of connection between the measuring point and the satellite is v r :

[0073]

[0074] Among them, (spx, spy, spz), (rpx, rpy, rpz) are the three-dimensional coordinates of the satellite and the measuring point; (svx, svy, svz), (rvx, rvy, rvz) are the three-dimensional velocities of the satellite and the measuring point;

[0075] The Doppler shift of the signal received by the satellite navigation signal simulator at the measurement point and the navigation satellite signal frequency are:

[0076] f d =-f c ·v r / c (15)

[0077] f=f c ·(1-v r / c) (16)

[0078] It can be seen from this that the Doppler shift is in the opposite direction to the relative motion, and the actual frequency is the frequency value after subtracting the Doppler shift;

[0079] 8) Pseudorange calculation: Pseudorange can be divided into two parts: the geometric distance between the satellite and the satellite navigation signal simulator; various error values ​​superimposed on the true distance; the pseudorange calculation formula is:

[0080]

[0081] in:

[0082] is the pseudorange generated from simulator k to satellite j;

[0083] is the geometric distance between simulator k and satellite j;

[0084] t k is the signal receiving time; t j is the time when the signal is transmitted;

[0085] δt j is the satellite clock error;

[0086] δt k is the simulator clock error;

[0087] is the ionospheric delay;

[0088] It is the tropospheric delay;

[0089] is the pseudorange measurement error caused by the multipath effect;

[0090] is the relativistic effect error;

[0091] It is the error caused by the satellite phase center deviating from the star's center of mass;

[0092] is the observation error;

[0093] The satellite navigation signal simulator is at sampling time t k When receiving satellite signals, the satellite position is Since it takes a period of time Δτ for a signal to propagate in space, the signal is at t j =t k - The satellite position at the time of transmission, which is transmitted from the satellite to the simulator at time Δτ That is, the signal is from The distance corresponding to the measured time of signal transmission from this position to the simulator is the pseudorange ρ:

[0094] ρ=cΔτ (18)

[0095] Where c is the speed of light, and its value can be taken as 2.99792458×10 8 m / s;

[0096] Since satellites are high-speed moving objects, their positions and There is a big difference. According to formula (17), the calculation of pseudorange Yes j The navigation satellite at time t k The geometric distance between the satellite navigation signal simulators at the time of positioning, t k is the known signal reception time, and the signal transmission time t j is unknown and can only be determined by t j =t k -Δτ is calculated, and Δτ needs to be calculated according to formula (18). It can be seen that the calculation equation (17) of the pseudorange is a transcendental equation and needs to be solved through iteration. At the same time, since the pseudorange ρ includes multiple error terms, the accuracy of the error terms is limited, which leads to errors in the calculation of the pseudorange. Therefore, the solution of the pseudorange needs to be iterated many times to obtain accurate Δτ, and then obtain accurate pseudorange ρ.

[0097] 9) Carrier phase delay calculation: The carrier phase value corresponding to a certain moment cannot be directly measured. The carrier phase observation value is the phase difference between the reference carrier signal generated by the satellite navigation signal simulator at a certain moment and the satellite carrier signal received at this time. Since the carrier phase and the pseudo code are on the same signal, the error term introduced by the signal during propagation is the same for the carrier phase and the pseudo range. Therefore, in the carrier phase calculation, the error term that needs to be considered is the same as that in the pseudo range calculation. The formula is:

[0098]

[0099] in:

[0100] is the carrier phase delay from simulator k to satellite j;

[0101] is the geometric distance between simulator k and satellite j;

[0102] t k is the signal receiving time; t j is the time when the signal is transmitted;

[0103] δt j is the satellite clock error;

[0104] δt k is the simulator clock error;

[0105] is the ionospheric delay;

[0106] It is the tropospheric delay;

[0107] is the pseudorange measurement error caused by the multipath effect;

[0108] is the relativistic effect error;

[0109] It is the error caused by the satellite phase center deviating from the star's center of mass;

[0110] is the observation error;

[0111] λ is the carrier wavelength, is the number of integral cycles of the carrier phase, It is the part of the carrier phase that is less than a full cycle;

[0112] 10) Calculate the frequency control word; when running in real time, the FPGA sends an interrupt to the DSP every 8ms. After the DSP enters the interrupt, it calculates the frequency control word of each channel and sends it to the FPGA, so that the FPGA accumulates it;

[0113] 11) Monitor and correct FPGA accumulation results; when FPGA sends an interrupt, DSP accurately calculates the current simulation time, obtains the navigation message count, ms count, code chip count and code phase, compares the counts and code phases corresponding to FPGA and DSP, and if the error exceeds the threshold, fine-tunes the frequency control word and corrects the error until the error is below the threshold;

[0114] 12) The GNSS signal simulation of the satellite navigation system is completed.

[0115] A method for controlling a ground laboratory remote sensing flight simulation platform of the present invention calculates the center of gravity of the flight simulation platform to meet the 15% MAC-29% MAC aircraft center of gravity limit requirement, thereby simulating the center of gravity position of a real aircraft, and the steps are:

[0116] 1) Determination of the coordinate system of the flight simulation platform:

[0117] The XOY coordinate system is selected as the coordinate system of the platform, where O is the origin of the coordinate system, and its position is 1 meter backward from the nose along the horizontal line of the aircraft structure; the X axis passes through the origin O along the horizontal line of the aircraft structure, and the backward direction is positive; the Y axis passes through the origin O, in the symmetry plane of the aircraft, and is perpendicular to the X axis, and the upward direction is positive;

[0118] 2) Flight simulation platform weighing and center of gravity position calculation:

[0119] Push the flight simulation platform onto the weighbridge, and make the front and main landing gear wheels stop at the center of each weighbridge. The placement of the weighbridge should be consistent with the position of the landing gear wheels. Record the readings of each scale and you will get:

[0120]

[0121] Where, P is the platform weight; P 1 When it is in horizontal state, the reading of the front landing gear wheel scale; P 2 is the sum of the readings of the main landing gear wheel scales when in horizontal state; L 1 is the X-direction distance between the front landing gear wheel and the origin of the coordinate system; L 2 is the X-direction distance between the main landing gear wheel and the origin of the coordinate system; X T is the X-distance between the center of gravity of the platform and the origin of the coordinate system; Y T is the Y-direction distance between the center of gravity of the platform and the origin of the coordinate system. Since the flight simulation platform has not been significantly modified, its theoretical value h is taken;

[0122] 3) According to the requirements of the remote sensing flight mission, simulate the state of increasing or decreasing the remote sensing equipment, onboard facilities, fuel, and crew load on the remote sensing flight simulation platform in the ground laboratory, and calculate the center of gravity position of the flight simulation platform under this state:

[0123]

[0124] Among them, X T ′ is the X-direction distance between the center of gravity of the platform and the origin of the coordinate system; p i is the weight of the added remote sensing equipment, onboard facilities, fuel, and crew load; i is the X-direction distance between the added remote sensing equipment, onboard facilities, fuel, and crew loads and the origin of the coordinate system; p k The weight of the remote sensing equipment, onboard facilities, fuel, and passenger load is reduced; k is the X-direction distance between the reduced loads of remote sensing equipment, onboard facilities, fuel, and passengers and the origin of the coordinate system; ∑p i It is the sum of the weight of the added remote sensing equipment, onboard facilities, fuel, and passenger load; ∑p k is the sum of the reduced weight of remote sensing equipment, onboard facilities, fuel, and passenger load, p i , p k , l i , l k All can be measured; Y T ′ is the Y-direction distance between the center of gravity of the platform and the origin of the coordinate system. Since the flight simulation platform has not been significantly modified, its theoretical value h′ is taken;

[0125] 4) Calculate the position of the aircraft's center of gravity on the mean aerodynamic chord:

[0126]

[0127] Among them, X O is the X-direction distance between the leading edge of the mean aerodynamic chord and the origin of the coordinate system; Y O is the Y-direction distance between the leading edge of the mean aerodynamic chord and the origin of the coordinate system; θ is the angle between the mean aerodynamic chord and the X-axis; X TO Y is the projection of the aircraft's center of gravity on the mean aerodynamic chord to the leading edge of the mean aerodynamic chord; TO It is the vertical distance from the center of gravity of the aircraft to the mean aerodynamic chord.

[0128] 5) Mean Aerodynamic Chord (MAC) Conversion: The center of gravity of the flight simulation platform needs to be converted into a percentage of the mean aerodynamic chord:

[0129]

[0130] in, is the mean aerodynamic chord percentage of the flight platform’s center of gravity, MAC is the mean aerodynamic chord, which is a known value;

[0131] 6) Check the calculation results of the center of gravity. If If the result meets the 15%MAC-29%MAC aircraft center of gravity limit requirement, the flight simulation platform successfully simulates the center of gravity state of the real aircraft. Otherwise, it is necessary to check and correct the combination of loads on the platform and recalculate the center of gravity position of the flight simulation platform.

[0132] Due to the adoption of the above technical solution, the advantages of the present invention are:

[0133] 1. The present invention builds a remote sensing flight simulation platform in the laboratory, on which a variety of remote sensing equipment can be installed, the compatibility of the equipment installation can be checked, some airworthiness tests can be conducted on the platform and the equipment, the functions of the equipment can be tested jointly, and remote sensing mission operation planning, remote sensing flight operation simulation tests, aircraft weight and center of gravity simulation tests, and other tasks can be completed.

[0134] 2. Problems that arise during the installation, power-on, and debugging of equipment can be discovered and solved on the laboratory platform to prevent similar problems from occurring when technicians travel to the airport for work.

[0135] 3. When a problem is found on the laboratory platform, multiple technicians can be called upon to participate in solving the problem, avoiding the problem of weak technical strength when technicians are working at the airport.

[0136] 4. The present invention includes a boarding ladder and a fuselage mounting bracket, has a stable structure, and is convenient and safe to load and unload equipment, and can ensure that multiple devices and multiple people can enter the platform at the same time.

[0137] 5. The remote sensing equipment needs to be in an open area outdoors to receive GNSS data. The present invention includes a satellite navigation signal simulator, which can provide GNSS data for the remote sensing equipment in the laboratory to ensure the normal progress of the experiment.

[0138] 6. Convenient and fast, can perform quantitative analysis, has high test accuracy, can improve work efficiency, and reduce repetitive work and work losses. BRIEF DESCRIPTION OF THE DRAWINGS

[0139] Figure 1 The following is a working flow chart of the satellite navigation signal simulator of the present invention;

[0140] Figure 2 This is a flow chart of the method for simulating the center of gravity state of a real aircraft in the present invention;

[0141] Figure 3 It is a schematic diagram of the overall structure of the present invention;

[0142] Figure 4 for Figure 3 Bottom view of

[0143] Figure 5 is an enlarged schematic diagram of the truss of the present invention;

[0144] Figure 6 It is an enlarged schematic diagram of a connecting member 1 of the present invention;

[0145] Figure 7 It is an enlarged schematic diagram of the second connecting member of the present invention;

[0146] Figure 8 It is an enlarged schematic diagram of the connector three of the present invention;

[0147] Fig. 9 It is an enlarged schematic diagram of the fourth connecting member of the present invention;

[0148] Fig.10 It is an enlarged schematic diagram of the stabilizing platform of the present invention.

[0149] In the figure, 1, fuselage; 2, left pod; 3, right pod; 4, left radome; 5, right radome; 6, downward optical window one; 7, downward optical window two; 8, downward optical window three; 9, truss; 10, non-airtight opening; 11, power distribution cabinet; 12, seat floor rail; 13, wiring box; 14, belly radome; 15, truss connecting support rod; 16, frame; 17, cross bar; 18, reinforcement rib; 19, connector one; 20, connector two; 21, connector three; 22, connector four; 23, U-shaped frame; 24, T-shaped plug plate; 25, bottom plate with socket; 26, branches; 27, vertical plate; 28, horizontal plate; 29, horizontal rod; 30, rectangular structure; 31, through hole; 32, handle; 33, cover plate. DETAILED DESCRIPTION

[0150] In order to make the objectives, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with specific embodiments and with reference to the accompanying drawings.

[0151] See attached Figure 1-Figure 10 A ground laboratory remote sensing flight simulation platform of the present invention comprises:

[0152] Fuselage, seat rail, floor, wiring box, non-airtight opening, downward optical window, optical glass, manual opening mechanism of camera door, manual opening mechanism of non-airtight protection door, satellite navigation signal simulator, power divider, pod, truss, radome, fuselage mounting bracket, boarding ladder, power system, platform center of gravity position; among which, seat rail, floor, wiring box, airtight cover of non-airtight opening, downward optical window, manual opening mechanism of camera door and non-airtight protection door, and satellite navigation signal simulator are installed inside the cabin; truss is connected to the bottom of fuselage, pod is connected to truss, and radome is installed on the outer surface of fuselage; fuselage mounting bracket is installed under fuselage structure, and boarding ladder is installed at the side door of fuselage; one end of power system is installed outside fuselage, and the other end is installed inside cabin;

[0153] The fuselage starts from the rear end of the cockpit and ends at the front end of the tail cabin, including the structure below the lower edge of the aircraft porthole, including the structure with the floor facing downward and the structure with the floor facing upward to the lower edge of the porthole;

[0154] The seat floor rail is installed inside the cabin and fixedly connected to the floor, and is used to install seats, cabinets, control consoles, and remote sensing equipment;

[0155] The floor is laid on the aircraft floor beam and is fixed to the floor beam by screws;

[0156] The wiring box is installed on the left and right sides and the middle position of the cabin. The wiring boxes on the left and right sides are fixedly connected to the side walls of the cabin, and the wiring box in the middle is fixedly connected to the floor. It is used for laying cables between mission equipment. The wiring box can prevent the cables from being stepped on and at the same time take into account the aesthetic requirements; openings are reserved on the inner sides of the wiring boxes on the left and right sides as channels for the mission equipment cables to enter the wiring boxes. The openings can be blocked with plugging covers, and a top cover is reserved on the top of the wiring box to facilitate opening and laying cable components;

[0157] The non-airtight opening is located in the tail cabin of the aircraft. Inside the airtight cover, a mounting platform parallel to the horizontal reference plane of the aircraft is arranged above the non-airtight opening. The mounting platform has mounting holes, and the remote sensing equipment and the airtight cover can be installed at the same time. The airtight cover plays an airtight role and is installed on the mounting platform to isolate the aircraft airtight cabin from the environment outside the cabin. The non-airtight opening is located at the outer skin of the belly of the aircraft, and the opening size is 700×700 mm measured in the flight direction×wing span direction.

[0158] There are three downward-looking optical windows located at the bottom of the cabin, and a stable platform for installing a camera is provided. The stable platform is a rectangular parallelepiped structure, and chamfers are provided at the four corners of the rectangular parallelepiped structure. A cover plate is provided on the top of the stable platform, and the cover plate is fixedly connected to the rectangular parallelepiped structure. A handle is provided on the cover plate, which can be used during installation and removal; a through hole is provided in the central part of the rectangular parallelepiped structure for the downward-looking optical channel of the camera;

[0159] The camera is mounted on a stable platform when in use, and the stable platform is mounted on a structural mounting platform;

[0160] The maximum weight that the structural mounting platform can bear is less than 150kg. The structural mounting platform is provided with four threaded holes for mounting corresponding to the four mounting holes of the stable platform. The geometric center axis of the optical glass in the downward optical window coincides with the hole axis of the stable platform. The distance between the mounting surface of the structural mounting platform and the upper surface of the optical glass should be less than 150mm. The parallelism between the mounting surface of the structural mounting platform and the upper surface of the optical glass is not greater than 1 degree. The mounting surface of the structural mounting platform is parallel to the horizontal plane of the aircraft.

[0161] There are three optical glasses, all of which are located on the three optical windows below the structural mounting platform. The effective aperture of the three optical glasses is ≥535mm×525mm measured in the front-to-back direction×the wingspan direction;

[0162] The manual opening mechanism of the camera cabin door is located below the cabin floor and is fixedly connected to the inner wall of the cabin bottom; the manual opening mechanism of the non-airtight protection door is located in the tail cabin of the aircraft and is fixedly connected to the tail cabin floor;

[0163] The camera door is located below the optical glass and is used to protect the camera lens and optical glass installed on the downward optical window, and to protect the camera lens and optical glass from being hit by gravel, ice, rain, snow or something brought up from the runway by the landing gear during takeoff and landing;

[0164] The non-airtight protection door is located below the non-airtight opening and close to the skin, and is used to protect the remote sensing equipment from damage during takeoff and landing of the aircraft; the non-airtight protection door is closed during takeoff and landing, and opened when the remote sensing equipment is operating; the non-airtight protection door can ensure the observation field angle of the mission equipment;

[0165] The satellite navigation signal simulator is fixedly connected to the side wall of the aircraft cabin; the power divider is fixedly connected to the side wall of the aircraft cabin; the power divider is used to provide satellite navigation signals for various remote sensing devices;

[0166] There are four trusses, two of which are installed on the left and right sides of the aircraft fuselage, with their upper ends fixedly connected to the side of the aircraft and their lower ends fixedly connected to the pod; the installation interface between the aircraft fuselage structure and the truss structure is connected by bolts;

[0167] The pods are installed on both sides of the fuselage of the aircraft through trusses, one on each side, with a center distance of 4.2 meters and an error of no more than 0.1 meters. The relative deformation of the two antennas installed in the pods is required to be less than 0.1 mm during stable flight;

[0168] There are three radomes in total, one radome is arranged under the belly of the fuselage, and a left radome and a right radome are arranged on both sides of the fuselage;

[0169] The function of the camera and lighting device: When the radome is closed, the movement of the antenna can be observed at any time, and a location can be selected for installation according to the aircraft conditions;

[0170] The fuselage mounting bracket is located below the fuselage structure and connected to the fuselage belly;

[0171] The boarding ladder is located at the boarding gate on the left side of the fuselage and is used for test personnel to board the aircraft. The load capacity is 300Kg;

[0172] The power supply system comprises a laboratory wall power supply, a power conversion device, a fuselage external grounding interface, a power distribution cabinet, and a power distribution box. The power conversion device is located near the laboratory wall power supply, the fuselage external grounding interface is installed on the fuselage structure surface, the power distribution cabinet is located at the rear end of the cabin cockpit, the power distribution box comprises a power distribution box 1 (P1), a power distribution box 2 (P2), and a power distribution box 3 (P3), the power distribution box 1 (P1), the power distribution box 2 (P2), and the power distribution box 3 (P3) are all located inside the cabin, the power distribution box 1 (P1) is located in the front equipment area, the power distribution box 2 (P2) is located in the rear equipment area, and the power distribution box 3 (P3) is located in the tail cabin equipment area, and the three distribution boxes are all fixedly connected to the side wall of the cabin;

[0173] The center of gravity position of the platform changes with the different installation positions of the equipment on the platform, and the center of gravity position of the platform in different installation states needs to be calculated; obtaining an accurate center of gravity position of the platform can ensure the smooth flight of the aircraft, thereby ensuring the quality of remote sensing data acquisition.

[0174] The location layout of the remote sensing equipment determines the center of gravity of the platform; after the remote sensing equipment acquires GNSS data, it is necessary to calculate the GNSS eccentricity component based on its own location and the location of the satellite navigation signal simulator for post-processing of the GNSS data.

[0175] Among them, the truss includes a frame, four truss connecting rods, two cross bars, reinforcement ribs, connector one, connector two, connector three, and connector four. The four truss connecting rods are hinged at the four corners of the frame. The four truss connecting rods are used to connect the pod. The two cross bars are fixed at the two ends of the frame. The two cross bars are components for lifting. The reinforcement ribs are fixed in the frame to increase the mechanical strength of the frame. Connector one, connector two, connector three, and connector four are used to connect the fuselage.

[0176] Among them, the connecting member 1 is provided with a U-shaped frame, a T-shaped plug plate, and a bottom plate with a socket, wherein the two ends of the U-shaped frame are connected to the frame, the bottom plate with the socket is connected to the frame, and the bottom plate with the socket is located in the U-shaped frame, the T-shaped plug plate can be inserted into the bottom plate with the socket, and can be connected to the bottom plate with the socket by bolts; a screw hole is provided on the top of the T-shaped plug plate, which can be used to connect to the fuselage; the connecting member 2 is provided with a branch and a vertical plate, the branch is connected to the middle of the frame, the bottom of the vertical plate is connected to the branch, and the vertical plate is provided with screw holes; the connecting member 3 is a horizontal plate, which is connected to the frame and has screw holes; the connecting member 4 is a transverse rod, and the transverse rod is provided with screw holes for connecting to the fuselage.

[0177] The pod includes a front fairing, a rear fairing and a middle section cabin, which are an integral structure. The middle section cabin includes an outer shell, an antenna cover and a connecting end plate, which are fixed on the outer shell, and both ends of the antenna cover are connected to the connecting end plate.

[0178] The functions of the camera and lighting device are: when the antenna cover is closed, the movement of the antenna can be observed at any time, and a location can be selected for installation according to the aircraft situation.

[0179] Among them, 6 mounting plates are arranged in the belly antenna cover, which are symmetrically distributed on both sides of the symmetry axis of the aircraft; the center distance between the left and right mounting plates is 600-800mm.

[0180] A control method for a ground laboratory remote sensing flight simulation platform of the present invention, wherein the satellite navigation signal simulator provides GNSS data for remote sensing equipment on the ground laboratory remote sensing flight simulation platform by simulating RF signals of multiple satellite navigation systems, and the steps are:

[0181] 1) DSP initialization, including reading simulation parameters, determining simulation time, calculating satellite navigation signal simulator trajectory, and reading satellite ephemeris parameters;

[0182] 2) Navigation message calculation: To simulate the navigation message, it is necessary to read the broadcast ephemeris file, which can be downloaded from the CDDIS website of the data center under IGS. The message parameters are extracted from the broadcast ephemeris file and converted into non-polar binary code or polar binary code. Then, the parity check code is generated by the check code encoder, converted into a relative code, and then modulo-two addition is performed with the folding code. Finally, a time mark is added to the end of each string to complete the encoding of the navigation message. Among them, the complete navigation message sequence contains 15,000 bits in a superframe, with a time length of 150 seconds. Each superframe consists of 5 frames, each frame consists of 3,000 bits, with a time length of 30 seconds. Each frame contains 15 strings, each string has a total of 200 bits, and the time length is 2 seconds.

[0183] 3) Satellite visibility judgment;

[0184] There is a relative motion between the satellite and the earth. After a period of time, there will inevitably be a state change of rising and falling stars at the same earth coordinate point. When the satellite's pitch angle is greater than 5 degrees, the satellite is considered to be visible; when the pitch angle is less than 5 degrees, the satellite is considered to be invisible. The broadcast ephemeris file generally stores the ephemeris parameters of nearly 30 navigation satellites in different time periods. After receiving the ephemeris parameters of all satellites, the DSP calculates the position of the satellite according to the set simulation start time, and then calculates the relative position of the two according to the initial position of the satellite navigation signal simulator, thereby obtaining the pitch angle and azimuth.

[0185] In the GPS system, the measurement coordinate system is established with the phase center of the GPS receiving antenna, and the elevation and azimuth of the satellite are calculated in this coordinate system. It is known that the position of the GPS satellite in the WGS84 coordinate system at time t is X tWGS-84 =(x T ,y T , z T ) T , the target position is X = (x, y, z) T ; The position of the target in the earth coordinate system can be calculated as (L, B, H) and the position of the satellite in the measurement coordinate system is The T in the footer refers to time T, and C refers to the satellite; the calculation formula for the satellite's elevation angle E and azimuth angle A is:

[0186]

[0187] in, If the z value of the target position is greater than or equal to 0, take 0; if z is less than 0, take π. A is an angle value.

[0188] In the GLONASS system, the satellite position is also converted from the geocentric coordinates to the station-centric coordinate system centered on the satellite navigation signal simulator position P;

[0189] Assume [X p Y p Z p ] T is the satellite station center coordinate, [X sv Y sv Z sv ] T and [X rec Y rec Z rec ] T are the coordinates of the satellite and the satellite navigation signal simulator in the PZ-90 geocentric coordinate system, B rec and L rec are the latitude and longitude of the satellite navigation signal simulator in the geodetic coordinate system, respectively.

[0190]

[0191] In the formula, H is

[0192]

[0193] After calculating the satellite's station center coordinates, the elevation angle E and azimuth angle A of the satellite navigation signal simulator observing the satellite can be obtained:

[0194]

[0195] If E ≥ E min , E min If it is the minimum elevation angle of the satellite, the satellite is a visible satellite. In subsequent processing, it is necessary to calculate the DOP value, error parameters, and signal status parameters for this satellite. Otherwise, it is an invisible satellite.

[0196] 4) Satellite position and speed calculation;

[0197] The equation describing the satellite motion trajectory can be obtained from the ICD file (space signal interface control file) of the satellite navigation system:

[0198]

[0199]

[0200] Where (x, y, z) represents the satellite coordinates, which is a function of time t; Vx(t), Vy(t), and Vz(t) are the first-order derivatives of the satellite coordinates with respect to time, representing the satellite's velocity; It is the second-order derivative of the coordinate with respect to time, indicating the acceleration of the satellite; the earth's gravitational constant μ = 398600440m 3 / s 2 ; Reference ellipsoid major radius Re =6378136m; Gravity potential second band harmonic coefficient The Earth's rotation speed ω=7.292115×10 -5 rad / s; It is the sum of the accelerations of the sun and the moon;

[0201] The fourth-order Runge-Kutta method is used to calculate the fourth slope value k under the premise of selecting the step size. 1 , k 2 , k 3 , k 4 , and use its weighted average as the average of the weighted slope, so as to obtain the approximate values ​​of the satellite position and velocity in the interval:

[0202]

[0203] Among them, k 1 is the slope at the beginning of the time period, k 2 is the slope at the midpoint of the time period, k 3 It is also the slope at the midpoint of the time period, k 4 is the slope at the end of the time period; y(i) is a vector function; x(i) is a vector; h is the integration step;

[0204] The number of integration times n required in the entire integration process is n=|t i -t b | / h, the integration time interval under normal circumstances |t i -t b | is not an integer multiple of the step size h, so the integration step size of the last step is smaller than the selected step size, which is h l =|t i -t b |-n*h, then after the entire integration process is completed, the integration time is the required solution time t i ;

[0205] Satellite signals are affected by the rotation of the earth during propagation. The satellite position and speed calculated by the above formula are the values ​​at the time of signal reception. In order to obtain the values ​​at the time of signal transmission, the result values ​​need to be corrected:

[0206]

[0207]

[0208] Among them, X(t), Y(t), Z(t) are satellite positions; U(t), V(t), W(t) are satellite velocities; Ω e is the average angular velocity of the Earth’s rotation; t is the end time of the integration interval; t eis the front moment of the integration interval; Ω e *(tt e ) is the angle that the earth rotates during the propagation of satellite signals;

[0209] 5) Calculation of ionospheric delay error: When GNSS signals pass through the ionosphere, electromagnetic wave refraction errors will occur, thereby reducing the accuracy and precision of satellite navigation positioning. According to the Klobuchar model, the ionospheric delay error can be corrected:

[0210]

[0211]

[0212] In the formula, α′ i =α i / 0.6, β′ i =β i (i=0, 1, 2, 3); parameter α 0 , α 1 , α 2 , α 3 , β 0 , β 1 , β 2 , β 3 is the ionospheric parameter in the ephemeris file; F is the tilt factor; φ m is the electromagnetic latitude of the intersection of the satellite signal propagation path and the central ionosphere; T p is the phase value of the cosine wave; P is the period of the cosine wave; t is the local time; E is the elevation angle of the satellite relative to the satellite navigation signal simulator;

[0213] 6) Calculation of tropospheric delay error: When the GNSS signal passes through the troposphere, the carrier, spread spectrum code and navigation message produce delay errors. The Hopfield model can be used to correct the delay error:

[0214]

[0215] In the formula, K d is the wet component delay in the zenith direction; K w is the dry component delay in the zenith direction; E is the elevation angle of the satellite relative to the satellite navigation signal simulator; η d , η w are the correction coefficients for the dry and wet components of the troposphere, respectively; T k and e are the absolute temperature and water vapor pressure at standard sea level respectively; h d is the upper limit of the tropospheric dry component based on sea level; h u is the simulator height; p 0 is the atmospheric pressure.

[0216] 7) Doppler shift calculation: There is a Doppler shift between the navigation satellite and the satellite navigation signal simulator. In order to restore the real satellite signal, the satellite navigation signal simulator needs to calculate the Doppler shift. According to the existing formula:

[0217]

[0218] Where v is the satellite speed, λ is the signal wavelength corresponding to the signal transmission frequency f, c is the speed of light, and β is the signal incident angle, that is, the angle between the receiving end and the signal incident direction; assuming that the signal frequency transmitted by the satellite is f c , the radial velocity in the direction of connection between the measuring point and the satellite is v r :

[0219]

[0220] Among them, (spx, spy, spz), (rpx, rpy, rpz) are the three-dimensional coordinates of the satellite and the measuring point; (svx, svy, svz), (rvx, rvy, rvz) are the three-dimensional velocities of the satellite and the measuring point;

[0221] The Doppler shift of the signal received by the satellite navigation signal simulator at the measurement point and the navigation satellite signal frequency are:

[0222] f d =-f c ·v r / c (15)

[0223] f=f c ·(1-v r / c) (16)

[0224] It can be seen from this that the Doppler shift is in the opposite direction to the relative motion, and the actual frequency is the frequency value after subtracting the Doppler shift;

[0225] 8) Pseudorange calculation: Pseudorange can be divided into two parts: the geometric distance between the satellite and the satellite navigation signal simulator; various error values ​​superimposed on the true distance; the pseudorange calculation formula is:

[0226]

[0227] in:

[0228] is the pseudorange generated from simulator k to satellite j;

[0229] is the geometric distance between simulator k and satellite j;

[0230] t k is the signal receiving time; tj is the time when the signal is transmitted;

[0231] δt j is the satellite clock error;

[0232] δt k is the simulator clock error;

[0233] is the ionospheric delay;

[0234] It is the tropospheric delay;

[0235] is the pseudorange measurement error caused by the multipath effect;

[0236] is the relativistic effect error;

[0237] It is the error caused by the satellite phase center deviating from the star's center of mass;

[0238] is the observation error;

[0239] The satellite navigation signal simulator is at sampling time t k When receiving satellite signals, the satellite position is Since it takes a period of time Δτ for a signal to propagate in space, the signal is at t j =t k - The satellite position at the time of transmission, which is transmitted from the satellite to the simulator at time Δτ That is, the signal is from The distance corresponding to the measured time of signal transmission from this position to the simulator is the pseudorange ρ:

[0240] ρ=cΔτ (18)

[0241] Where c is the speed of light, and its value can be taken as 2.99792458×10 8 m / s;

[0242] Since satellites are high-speed moving objects, their positions and There is a big difference. According to formula (17), the calculation of pseudorange Yes j The navigation satellite at time t k The geometric distance between the satellite navigation signal simulators at the time of positioning, t k is the known signal reception time, and the signal transmission time t j is unknown and can only be determined by t j =t k-Δτ is calculated, and Δτ needs to be calculated according to formula (18). It can be seen that the calculation equation (17) of the pseudorange is a transcendental equation and needs to be solved through iteration. At the same time, since the pseudorange ρ includes multiple error terms, the accuracy of the error terms is limited, which leads to errors in the calculation of the pseudorange. Therefore, the solution of the pseudorange needs to be iterated many times to obtain accurate Δτ, and then obtain accurate pseudorange ρ.

[0243] 9) Carrier phase delay calculation: The carrier phase value corresponding to a certain moment cannot be directly measured. The carrier phase observation value is the phase difference between the reference carrier signal generated by the satellite navigation signal simulator at a certain moment and the satellite carrier signal received at this time. Since the carrier phase and the pseudo code are on the same signal, the error term introduced by the signal during propagation is the same for the carrier phase and the pseudo range. Therefore, in the carrier phase calculation, the error term that needs to be considered is the same as that in the pseudo range calculation. The formula is:

[0244]

[0245] in:

[0246] is the carrier phase delay from simulator k to satellite j;

[0247] is the geometric distance between simulator k and satellite j;

[0248] t k is the signal receiving time; t j is the time when the signal is transmitted;

[0249] δt j is the satellite clock error;

[0250] δt k is the simulator clock error;

[0251] is the ionospheric delay;

[0252] It is the tropospheric delay;

[0253] is the pseudorange measurement error caused by the multipath effect;

[0254] is the relativistic effect error;

[0255] It is the error caused by the satellite phase center deviating from the star's center of mass;

[0256] is the observation error;

[0257] λ is the carrier wavelength, is the number of integral cycles of the carrier phase, It is the part of the carrier phase that is less than a full cycle;

[0258] 10) Calculate the frequency control word; when running in real time, the FPGA sends an interrupt to the DSP every 8ms. After the DSP enters the interrupt, it calculates the frequency control word of each channel and sends it to the FPGA, so that the FPGA accumulates it;

[0259] 11) Monitor and correct FPGA accumulation results; when FPGA sends an interrupt, DSP accurately calculates the current simulation time, obtains the navigation message count, ms count, code chip count and code phase, compares the counts and code phases corresponding to FPGA and DSP, and if the error exceeds the threshold, fine-tunes the frequency control word and corrects the error until the error is below the threshold;

[0260] 12) The GNSS signal simulation of the satellite navigation system is completed.

[0261] A method for controlling a ground laboratory remote sensing flight simulation platform of the present invention calculates the center of gravity of the flight simulation platform to meet the 15% MAC-29% MAC aircraft center of gravity limit requirement, thereby simulating the center of gravity position of a real aircraft, and the steps are:

[0262] 1) Determination of the coordinate system of the flight simulation platform:

[0263] The XOY coordinate system is selected as the coordinate system of the platform, where O is the origin of the coordinate system, and its position is 1 meter backward from the nose along the horizontal line of the aircraft structure; the X axis passes through the origin O along the horizontal line of the aircraft structure, and the backward direction is positive; the Y axis passes through the origin O, in the symmetry plane of the aircraft, and is perpendicular to the X axis, and the upward direction is positive;

[0264] 2) Flight simulation platform weighing and center of gravity position calculation:

[0265] Push the flight simulation platform onto the weighbridge, and make the front and main landing gear wheels stop at the center of each weighbridge. The placement of the weighbridge should be consistent with the position of the landing gear wheels. Record the readings of each scale and you will get:

[0266]

[0267] Where, P is the platform weight; P 1 When it is in horizontal state, the reading of the front landing gear wheel scale; P 2 is the sum of the readings of the main landing gear wheel scales when in horizontal state; L 1 is the X-direction distance between the front landing gear wheel and the origin of the coordinate system; L 2 is the X-direction distance between the main landing gear wheel and the origin of the coordinate system; X T is the X-distance between the center of gravity of the platform and the origin of the coordinate system; Y Tis the Y-direction distance between the center of gravity of the platform and the origin of the coordinate system. Since the flight simulation platform has not been significantly modified, its theoretical value h is taken;

[0268] 3) According to the requirements of the remote sensing flight mission, simulate the state of increasing or decreasing the remote sensing equipment, onboard facilities, fuel, and crew load on the remote sensing flight simulation platform in the ground laboratory, and calculate the center of gravity position of the flight simulation platform under this state:

[0269]

[0270] Among them, X T ′ is the X-direction distance between the center of gravity of the platform and the origin of the coordinate system; p i is the weight of the added remote sensing equipment, onboard facilities, fuel, and crew load; i is the X-direction distance between the added remote sensing equipment, onboard facilities, fuel, and crew loads and the origin of the coordinate system; p k The weight of the remote sensing equipment, onboard facilities, fuel, and passenger load is reduced; k is the X-direction distance between the reduced loads of remote sensing equipment, onboard facilities, fuel, and passengers and the origin of the coordinate system; ∑p i It is the sum of the weight of the added remote sensing equipment, onboard facilities, fuel, and passenger load; ∑p k is the sum of the reduced weight of remote sensing equipment, onboard facilities, fuel, and passenger load, p i , p k , l i , l k All can be measured; Y T ′ is the Y-direction distance between the center of gravity of the platform and the origin of the coordinate system. Since the flight simulation platform has not been significantly modified, its theoretical value h′ is taken;

[0271] 4) Calculate the position of the aircraft's center of gravity on the mean aerodynamic chord:

[0272]

[0273] Among them, X O is the X-direction distance between the leading edge of the mean aerodynamic chord and the origin of the coordinate system; Y O is the Y-direction distance between the leading edge of the mean aerodynamic chord and the origin of the coordinate system; θ is the angle between the mean aerodynamic chord and the X-axis; X TO Y is the projection of the aircraft's center of gravity on the mean aerodynamic chord to the leading edge of the mean aerodynamic chord; TO It is the vertical distance from the center of gravity of the aircraft to the mean aerodynamic chord.

[0274] 5) Mean Aerodynamic Chord (MAC) Conversion: The center of gravity of the flight simulation platform needs to be converted into a percentage of the mean aerodynamic chord:

[0275]

[0276] in, is the mean aerodynamic chord percentage of the flight platform’s center of gravity, MAC is the mean aerodynamic chord, which is a known value;

[0277] 6) Check the calculation results of the center of gravity. If If the result meets the 15%MAC-29%MAC aircraft center of gravity limit requirement, the flight simulation platform successfully simulates the center of gravity state of the real aircraft. Otherwise, it is necessary to check and correct the combination of loads on the platform and recalculate the center of gravity position of the flight simulation platform.

[0278] The laboratory wall power interface of the platform of the present invention is connected to the power conversion device through a power cord, the laboratory wall power interface is connected to the external grounding interface of the fuselage through a power cord, the power conversion device is connected to the distribution cabinet through the power cord, the distribution cabinet is respectively connected to the distribution box 1 P1, the distribution box 2 P2, and the distribution box 3 P3 through a DC power cord, and the power conversion device is connected to the distribution box 1 P1 and the distribution box 2 P2 through an AC power cord.

[0279] The power conversion device is a device that converts AC380V / 50Hz three-phase alternating current into AC115V / 400Hz alternating current. The power conversion device has two outputs. The first circuit simulates the circuit output from the aircraft generator to the distribution cabinet, and the second circuit is a circuit output through the distribution cabinet. The second circuit is an independent secondary circuit equipped for the mission system, independent of the aircraft simulation circuit. If a power supply failure occurs in the aircraft simulation circuit, the second circuit can be disconnected first; the total power of the power conversion device is 30kVA, of which the AC115V / 400Hz three-phase alternating current mission load is ≤5kVA, the AC115V / 50Hz single-phase alternating current mission load is ≤5kVA, and the 28V DC load is ≤20kW.

[0280] The power distribution cabinet is provided with a power supply and distribution connection port, and multiple power supplies can be drawn from the power distribution cabinet to provide power to the distribution box 1 P1, the distribution box 2 P2, and the distribution box 3 P3. The power distribution cabinet outputs 2 branches, branch 1 provides 28V DC power, and branch 2 provides AC115V / 400Hz three-phase AC power.

[0281] The distribution box P1 outputs 5 branches, branches 1, 2, and 3 of distribution box P1 provide 28V DC power, branch 4 of distribution box P1 provides AC115V / 400Hz three-phase AC power, and branch 5 of distribution box P1 provides AC115V / 400Hz single-phase AC power; the maximum power supply capacity of 28V DC power is 15kW, the maximum power supply capacity of AC115V / 400Hz three-phase AC power is 5kVA, and the maximum power supply capacity of AC115V / 50Hz single-phase AC power is 0.5kVA.

[0282] The distribution box P2 outputs 6 branches, branch 1 of the distribution box P2 provides AC115V / 400Hz three-phase AC power, branches 2 and 6 of the distribution box P2 provide AC115V / 400Hz single-phase AC power, and branches 3, 4, and 5 of the distribution box P2 provide 28V DC power; the maximum power supply capacity of 28V DC power is 8kW, the maximum power supply capacity of AC115V / 400Hz three-phase AC power is 5kVA, and the maximum power supply capacity of AC115V / 50Hz single-phase AC power is 5kVA.

[0283] The distribution box three P3 outputs one branch, and branch 1 of the distribution box three P3 provides 28V DC power; the maximum power supply capacity of 28V DC power is 5kW.

[0284] The output DC voltage of the distribution cabinet is 24.8V-30.8V; under full load, the line voltage drop from the distribution cabinet to distribution box one P1 is 0.4V; the line voltage drop from the distribution cabinet to distribution box two P2 is 0.9V; the line voltage drop from the distribution cabinet to distribution box three P3 is 1.4V.

[0285] The present invention builds a remote sensing flight simulation platform, and can complete the installation, airworthiness test, joint debugging test, mission planning, simulated flight and other tasks of various remote sensing equipment in the laboratory.

[0286] Remote sensing equipment: including equipment with different functions such as high-resolution linear array digital aerial cameras, high-resolution area array digital aerial cameras, multi-modal digital aerial cameras, wide-band imaging spectrometers, push-broom hyperspectral imagers, three-dimensional lidars, high-resolution polarimetric interferometric synthetic aperture radars, multi-band fully polarimetric synthetic aperture radars, fully polarimetric microwave radiometers / scatterometers, atmospheric environment lidars, differential optical absorption spectrometers, multi-angle polarimetric radiometers, etc.

[0287] Power divider: a device that divides one input signal energy into two or more outputs with equal or unequal energy. The satellite navigation signal of the satellite navigation signal simulator is led to the input end of the power divider through a coaxial cable, and then the output end of the power divider is connected to a variety of remote sensing equipment through multiple coaxial cables to provide satellite navigation signals for a variety of remote sensing equipment.

[0288] Checksum encoder; an encoding algorithm used to generate a checksum for data.

[0289] Relative code: A type of coding in communication principles. In relative codes, "1" and "0" are represented by level jumps or unchanged levels respectively; if level jumps are used to represent "1", it is called a transmission-relative code (in telegraph communication, "1" is often called a transmission number, and "0" is called a space number); if level jumps are used to represent "0", it is called a space number relative code.

[0290] Folding code: In the folding code, the first bit on the left represents the positive or negative sign (signal polarity), and the second bit to the last bit represent the signal amplitude; the first bit is 1 for positive and 0 for negative; the folding codes with the same absolute value have the same code group except the first bit, and are symmetrically folded relative to the zero level (between the 7th level and the 8th level), so this code group is called a folding code.

[0291] Modulo 2 addition: The so-called "modulo 2 addition" is the addition between 0 and 1, where 0+0=0, 1+0=0+1=1, and 1+1=0(!). This operation is commonly used in communications and computers. You can think of 0 and 1 as "even" and "odd", respectively. Then the first two equations represent: even plus even equals even, odd plus even equals odd, and 1+1=0 means odd plus odd equals even. For any number a 1 , a 2 , …, a m (each is 0 or 1), we can add them modulo 2 1 +a 2 +…+a m ; It is not difficult to see that when there are an odd number of 1s in these m numbers, the result is 1, otherwise the result is 0.

[0292] IGS: International GNSS Service (IGS) is an international cooperation organization established in 1993 by the International Association of Geodesy (IAG) to support geodesy and geodynamic research. It officially started work on January 1, 1994. Experiments such as the Global GPS Campaign from June to September 1992 laid the foundation for the establishment of IGS. Since then, with the completion and commissioning of other global satellite navigation and positioning systems such as GLONASS, the International GPS Service has also expanded its scope of work and was renamed the International GNSS Service (International GNSS Service), but it is still abbreviated as IGS.

[0293] The CDDIS website, one of the global data centers under IGS, provides free pseudorange and phase observations, broadcast ephemeris, meteorological and other measured data.

[0294] Navigation message: A navigation message is a message broadcast by a navigation satellite to users that describes the operating status parameters of the navigation satellite, including system time, ephemeris, almanac, correction parameters of the satellite clock, navigation satellite health status and ionospheric delay model parameters. The parameters of the navigation message provide users with time information, and the user's position coordinates and speed can be calculated using the navigation message parameters.

[0295] Broadcast ephemeris file, satellite orbit parameter telegram information within a certain period of time predicted by satellite broadcast; also called forecast ephemeris; the content of forecast ephemeris usually includes Kepler orbit parameters and necessary orbit perturbation correction parameters relative to a certain reference epoch; the satellite Kepler orbit parameters of the corresponding reference epoch are also called reference ephemeris; the reference ephemeris only represents the orbit parameters of the satellite at the reference epoch, but under the influence of perturbation, the actual orbit of the satellite may deviate from the reference orbit; the degree of deviation mainly depends on the time difference between the observed epoch and the selected reference epoch; if the known satellite reference ephemeris is corrected by the perturbation term of the orbit parameter, the satellite ephemeris of any observed epoch can be extrapolated; the selection of broadcast ephemeris parameters adopts the method of Kepler orbit parameters plus harmonic term correction. There are 16 GPS broadcast ephemeris parameters: 1 reference time parameter, 6 Kepler orbit parameters corresponding to the reference time and 9 parameters reflecting the influence of perturbation.

[0296] Satellite ephemeris parameters, there are 16 ephemeris parameters in total, including 1 reference time, 6 Kepler orbit parameters corresponding to the reference time and 9 parameters reflecting the influence of perturbations.

[0297] Message parameters; including satellite ephemeris, satellite clock error, ionospheric delay correction parameters, almanac data and time synchronization parameters, etc.

[0298] DSP: DSP (digital signal processor) is a unique microprocessor that processes large amounts of information using digital signals. Its working principle is to receive analog signals, convert them into digital signals of 0 or 1, and then modify, delete, and enhance the digital signals, and interpret the digital data back to analog data or actual environment formats in other system chips; it is not only programmable, but also has a real-time running speed of tens of millions of complex instruction programs per second, far exceeding general-purpose microprocessors, and is an increasingly important computer chip in the digital electronic world. Its powerful data processing capabilities and high operating speed are the two most commendable features.

[0299] WGS84 coordinate system: an internationally adopted geocentric coordinate system; the coordinate origin is the center of mass of the earth, the Z axis of its geocentric rectangular coordinate system points to the direction of the agreed earth pole (CTP) defined by BIH (International Time Service) 1984.0, the X axis points to the intersection of the zero meridian plane of BIH 1984.0 and the CTP equator, and the Y axis is perpendicular to the Z axis and the X axis to form a right-handed coordinate system, called the 1984 World Geodetic Coordinate System.

[0300] GLONASS system: A second-generation military satellite navigation system independently developed and controlled by the Ministry of Defense of the Soviet Union (now Russia). Similar to the US GPS, this system also has a civilian window. GLONASS technology can provide high-precision three-dimensional position, three-dimensional velocity and time information for various military and civilian users in the world's land, sea, air and near-Earth space around the clock and continuously. GLONASS is superior to GPS after applying Selective Availability (SA) in terms of positioning, velocity measurement and timing accuracy, because Russia has promised to ICAO and Maritime Organization to provide civil navigation services to users around the world.

[0301] P2-90 Geocentric Coordinate System: The coordinate system adopted by Russia's GLONASS navigation system in 1993.

[0302] ICD file: Navigation satellite interface control file.

[0303] Fourth-order Runge-Kutta: In numerical analysis, Runge-Kutta methods are an important class of implicit or explicit iterative methods for the solution of nonlinear ordinary differential equations. These techniques were invented by mathematicians Carl Runge and Martin Wilhelm Kutta around 1900; Runge-Kutta methods are a high-precision single-step algorithm widely used in engineering, including the famous Euler method, for numerically solving differential equations. Due to the high accuracy of this algorithm and the measures taken to suppress errors, its implementation principle is also relatively complex; the classic Runge-Kutta method is fourth-order, that is, the weighted average of the slopes at four points is used as an approximation of the average slope, forming a series of fourth-order Runge-Kutta formulas with fourth-order accuracy.

[0304] GNSS signal: Global Navigation Satellite System signal.

[0305] Klobuchar model: A method of ionospheric delay correction suitable for GPS single-frequency receivers proposed by American scientist Klobuchar in 1987.

[0306] Hopfield model: tropospheric delay error model.

[0307] FPGA: Field Programmable Gate Array, emerged as a semi-custom circuit in the field of Application Specific Integrated Circuit (ASIC), which not only solves the shortcomings of custom circuits, but also overcomes the disadvantage of the limited number of gate circuits of the original programmable devices.

[0308] RF signal: radio frequency signal.

[0309] MAC: Mean aerodynamic chord. The mean aerodynamic chord is the longitudinal characteristic length of the aircraft and is a particularly important geometric parameter. For an actual wing of any plane shape, its chord length varies from the wing root to the wing tip. We can imagine that there is a fairly rectangular wing with the same area as the actual wing, and the same pitch moment and aerodynamic force. We call the chord of such a rectangular wing the mean aerodynamic chord of the wing.

[0310] Obviously, the above embodiments of the present invention are merely examples for clearly explaining the present invention, and are not intended to limit the implementation methods of the present invention. For those skilled in the art, other different forms of changes or modifications can be made based on the above description. It is impossible to list all implementation methods here. Any obvious changes or modifications derived from the technical solution of the present invention are still within the scope of protection of the present invention.

Claims

1. A ground laboratory remote sensing flight simulation platform, characterized in that: include: Fuselage, seat rail, floor, wiring box, non-airtight opening, downward optical window, optical glass, manual opening mechanism of camera door, manual opening mechanism of non-airtight protection door, satellite navigation signal simulator, power divider, pod, truss, radome, fuselage mounting bracket, boarding ladder, power system, platform center of gravity position; among which, seat rail, floor, wiring box, airtight cover of non-airtight opening, downward optical window, manual opening mechanism of camera door and non-airtight protection door, and satellite navigation signal simulator are installed inside the cabin; truss is connected to the bottom of fuselage, pod is connected to truss, and radome is installed on the outer surface of fuselage; fuselage mounting bracket is installed under fuselage structure, and boarding ladder is installed at the side door of fuselage; one end of power system is installed outside fuselage, and the other end is installed inside cabin; The fuselage starts from the rear end of the cockpit and ends at the front end of the tail cabin, including the structure below the lower edge of the aircraft porthole, including the structure with the floor facing downward and the structure with the floor facing upward to the lower edge of the porthole; The seat floor rail is installed inside the cabin and fixedly connected to the floor, and is used to install seats, cabinets, control consoles, and remote sensing equipment; The floor is laid on the aircraft floor beam and is fixed to the floor beam by screws; The wiring boxes are installed on the left and right sides and the middle of the cabin. The wiring boxes on the left and right sides are fixedly connected to the side walls of the cabin, and the wiring box in the middle is fixedly connected to the floor, and are used for laying cables between remote sensing equipment; openings are reserved on the inner sides of the wiring boxes on the left and right sides as channels for mission equipment cables to enter the wiring boxes, and the openings can be blocked with plugging covers. A top cover is reserved on the top of the wiring box to facilitate opening and laying of cable components; The non-airtight opening is located in the tail cabin of the aircraft. Inside the airtight cover, a mounting platform parallel to the horizontal reference plane of the aircraft is arranged above the non-airtight opening. The mounting platform has mounting holes, and the remote sensing equipment and the airtight cover can be installed at the same time. The airtight cover plays an airtight role and is installed on the mounting platform to isolate the aircraft airtight cabin from the environment outside the cabin. The non-airtight opening is located at the outer skin of the belly of the aircraft, and the opening size is 700×700 mm measured in the flight direction×wing span direction. There are three downward-looking optical windows located at the bottom of the cabin, and a stable platform for installing a camera is provided. The stable platform is a rectangular parallelepiped structure, and chamfers are provided at the four corners of the rectangular parallelepiped structure. A cover plate is provided on the top of the stable platform, and the cover plate is fixedly connected to the rectangular parallelepiped structure. A handle is provided on the cover plate, which can be used for installation and removal; a through hole is provided in the central part of the rectangular parallelepiped structure for the downward-looking optical channel of the camera; The camera is mounted on a stable platform when in use, and the stable platform is mounted on a structural mounting platform; The maximum weight that the structural mounting platform can bear is less than 150kg. The structural mounting platform is provided with four threaded holes for mounting corresponding to the four mounting holes of the stable platform. The geometric center axis of the optical glass in the downward optical window coincides with the hole axis of the stable platform. The distance between the mounting surface of the structural mounting platform and the upper surface of the optical glass should be less than 150mm. The parallelism between the mounting surface of the structural mounting platform and the upper surface of the optical glass is not greater than 1 degree. The mounting surface of the structural mounting platform is parallel to the horizontal plane of the aircraft. There are three optical glasses, all of which are located on the three optical windows below the structural mounting platform. The effective aperture of the three optical glasses is ≥535mm×525mm measured in the front-to-back direction×the wingspan direction; The manual opening mechanism of the camera cabin door is located below the cabin floor and is fixedly connected to the inner wall of the cabin bottom; the manual opening mechanism of the non-airtight protection door is located in the tail cabin of the aircraft and is fixedly connected to the tail cabin floor; The photographic cabin door is located below the optical glass; The non-airtight protection door is located below the non-airtight opening and close to the skin; the non-airtight protection door is closed during takeoff and landing, and opened when the remote sensing equipment is operating; the non-airtight protection door can ensure the observation field angle of the mission equipment; The satellite navigation signal simulator is fixedly connected to the side wall of the aircraft cabin; the power divider is fixedly connected to the side wall of the aircraft cabin; There are four trusses in total, two of which are installed on the left and right sides of the aircraft fuselage, and the upper ends thereof are fixedly connected to the side of the aircraft, and the lower ends are fixedly connected to the pod; the installation interface between the aircraft body structure and the truss structure adopts a bolt connection method; the truss includes a frame, four truss connecting rods, two cross bars, reinforcing ribs, connecting piece 1, connecting piece 2, connecting piece 3, and connecting piece 4; the four truss connecting rods are hinged at the four corners of the frame; the four truss connecting rods are used to connect the pod, the two cross bars are fixed at the two ends of the frame, the two cross bars are components for hoisting, the reinforcing ribs are fixed in the frame to increase the mechanical strength of the frame, and the connecting piece 1, connecting piece 2, connecting piece 3, and connecting piece 4 are used to connect the fuselage; The pods are installed on both sides of the fuselage of the aircraft through trusses, one on each side, with a center distance of 4.2 meters and an error of no more than 0.1 meters. The relative deformation of the two antennas installed in the pods is required to be less than 0.1 mm during stable flight; There are three radomes, one of which is arranged under the belly of the aircraft, and a left radome and a right radome are arranged on both sides of the fuselage; six mounting plates are arranged in the belly radome, symmetrically distributed on both sides of the symmetry axis of the aircraft; the center distance between the left and right mounting plates is 600-800 mm; The function of the camera and lighting device: When the radome is closed, the movement of the antenna can be observed at any time, and a location can be selected for installation according to the aircraft situation; The fuselage mounting bracket is located below the fuselage structure and connected to the fuselage belly; The boarding ladder is located at the boarding gate on the left side of the fuselage and is used for test personnel to board the aircraft. The load capacity is 300Kg; The power supply system comprises a laboratory wall power supply, a power conversion device, a fuselage external grounding interface, a power distribution cabinet, and a power distribution box. The power conversion device is located near the laboratory wall power supply. The fuselage external grounding interface is installed on the fuselage structure surface. The power distribution cabinet is located at the rear end of the cabin cockpit. The power distribution box comprises a power distribution box 1 (P1), a power distribution box 2 (P2), and a power distribution box 3 (P3). The power distribution box 1 (P1), the power distribution box 2 (P2), and the power distribution box 3 (P3) are all located inside the cabin. The power distribution box 1 (P1) is located in the front equipment area, the power distribution box 2 (P2) is located in the rear equipment area, and the power distribution box 3 (P3) is located in the tail cabin equipment area. The three power distribution boxes are all fixedly connected to the side wall of the cabin. The center of gravity position of the platform changes with the different installation positions of the equipment on the platform, and the center of gravity position of the platform in different installation states needs to be calculated; The location layout of the remote sensing equipment determines the center of gravity of the platform; after the remote sensing equipment acquires GNSS data, it is necessary to calculate the GNSS eccentricity component based on its own location and the location of the satellite navigation signal simulator for post-processing of the GNSS data.

2. A ground laboratory remote sensing flight simulation platform as claimed in claim 1, characterized in that: The first connector is provided with a U-shaped frame, a T-shaped plug plate, and a bottom plate with a socket, wherein both ends of the U-shaped frame are connected to the frame, the bottom plate with the socket is connected to the frame, and the bottom plate with the socket is located in the U-shaped frame, the T-shaped plug plate can be inserted into the bottom plate with the socket, and can be connected to the bottom plate with the socket by bolts; a screw hole is provided on the top of the T-shaped plug plate, which can be used to connect to the fuselage; the second connector is provided with a branch and a vertical plate, the branch is connected to the middle of the frame, the bottom of the vertical plate is connected to the branch, and the vertical plate is provided with a screw hole; the third connector is a horizontal plate, which is connected to the frame and has a screw hole; the fourth connector is a transverse rod, and the transverse rod is provided with a screw hole for connecting to the fuselage.

3. A ground laboratory remote sensing flight simulation platform as claimed in claim 1, characterized in that: The pod includes a front fairing, a rear fairing, and a middle section cabin. The front fairing, the rear fairing, and the middle section cabin are an integral structure. The middle section cabin includes an outer shell, an antenna cover, and a connecting end plate. The connecting end plate is fixed on the outer shell, and both ends of the antenna cover are connected to the connecting end plate.

4. A ground laboratory remote sensing flight simulation platform as claimed in claim 1, characterized in that: The function of the camera and lighting device is to observe the movement of the antenna at any time when the antenna cover is closed, and select a location for installation according to the aircraft situation.

5. A control method for the ground laboratory remote sensing flight simulation platform according to claim 1, characterized in that: The satellite navigation signal simulator provides GNSS data for remote sensing equipment on a remote sensing flight simulation platform in a ground laboratory by simulating RF signals of multiple satellite navigation systems. The steps are: 1) DSP initialization, including reading simulation parameters, determining simulation time, calculating satellite navigation signal simulator trajectory, and reading satellite ephemeris parameters; 2) Navigation message calculation: To simulate the navigation message, it is necessary to read the broadcast ephemeris file, which can be downloaded from the CDDIS website of the data center under IGS. The message parameters are extracted from the broadcast ephemeris file and converted into non-polar binary code or polar binary code. Then, the parity check code is generated by the check code encoder, converted into a relative code, and then modulo-2 addition is performed with the folding code. Finally, a time mark is added at the end of each string to complete the encoding of the navigation message. 3) Satellite visibility judgment; There is a relative motion between the satellite and the earth. After a period of time, there will inevitably be a state change of rising and falling stars at the same earth coordinate point. When the satellite's pitch angle is greater than 5 degrees, the satellite is considered to be visible; when the pitch angle is less than 5 degrees, the satellite is considered to be invisible. The broadcast ephemeris file generally stores the ephemeris parameters of nearly 30 navigation satellites in different time periods. After receiving the ephemeris parameters of all satellites, the DSP calculates the position of the satellite according to the set simulation start time, and then calculates the relative position of the two according to the initial position of the satellite navigation signal simulator, thereby obtaining the pitch angle and azimuth. In the GPS system, a measurement coordinate system is established with the phase center of the GPS receiving antenna, and the elevation and azimuth of the satellite are calculated in this coordinate system; The position of the GPS satellite in the WGS84 coordinate system at this moment is , the target location is ; The position of the target in the geodetic coordinate system can be calculated as and the position of the satellite in the measurement coordinate system is , the T in the footnote refers to the time T, and C refers to the satellite; the elevation angle of the satellite and azimuth The calculation formula is: （1） in, ; In the GLONASS system, the satellite position is also converted from the geocentric coordinates to the station-centric coordinate system centered on the satellite navigation signal simulator position P; set up is the satellite station center coordinate, and are the coordinates of the satellite and the satellite navigation signal simulator in the PZ-90 geocentric coordinate system, and are the latitude and longitude of the satellite navigation signal simulator in the geodetic coordinate system, respectively. （2） In the formula, for （3） After calculating the satellite's station center coordinates, the elevation angle of the satellite navigation signal simulator observing the satellite can be calculated. and azimuth : （4） if , If it is the minimum elevation angle of the satellite, the satellite is a visible satellite. In subsequent processing, the DOP value, error parameters, and signal status parameters need to be calculated for this satellite. Otherwise, it is an invisible satellite. 4) Satellite position and speed calculation; The equation describing the satellite motion trajectory can be obtained from the ICD file of the satellite navigation system: （5） （6） in, represents the satellite coordinates, which is the time Function of is the first derivative of the satellite coordinates with respect to time, indicating the velocity of the satellite; is the second-order derivative of the satellite coordinates with respect to time, indicating the acceleration of the satellite; the Earth's gravitational constant ; Reference ellipsoid major radius ; Gravity potential second band harmonic coefficient ; Earth's rotation speed ; It is the sum of the accelerations of the sun and the moon; ; The fourth-order Runge-Kutta method is used to calculate the fourth slope value under the premise of selecting the step size. , and use its weighted average as the average of the weighted slope, so as to obtain the approximate values ​​of the satellite position and velocity in the interval: （7） in, is the slope at the beginning of the time period, is the slope at the midpoint of the time period, is also the slope at the midpoint of the time period, is the slope at the end of the time period; is a vector function; is a vector; is the integration step size; The number of points required in the entire integration process , the integration time interval under normal circumstances is Not step length An integer multiple of , so the integration step size of the last step is smaller than the selected step size, which is , then after the entire integration process is completed, the integration time is the required solution time ; Satellite signals are affected by the rotation of the earth during propagation. The satellite position and speed calculated by the above formula are the values ​​at the time of signal reception. In order to obtain the values ​​at the time of signal transmission, the result values ​​need to be corrected: （8） (9) in, is the satellite position; is the satellite speed; is the average angular velocity of the Earth's rotation; is the back end moment of the integration interval; is the front moment of the integration interval; It is the angle that the earth rotates during the propagation of satellite signals; 5) Calculation of ionospheric delay error: When GNSS signals pass through the ionosphere, electromagnetic wave refraction errors will occur, thereby reducing the accuracy and precision of satellite navigation positioning. According to the Klobuchar model, the ionospheric delay error can be corrected: （10） （11） In the formula, ;parameter are the ionospheric parameters in the ephemeris file; is the tilt factor; It is the electromagnetic latitude of the intersection of the satellite signal propagation path and the central ionosphere; is the phase value of the cosine wave; is the period of the cosine wave; It is local time; is the elevation angle of the satellite relative to the satellite navigation signal simulator; 6) Calculation of tropospheric delay error: When the GNSS signal passes through the troposphere, the carrier, spread spectrum code and navigation message produce delay errors. The Hopfield model can be used to correct the delay error: （12） In the formula, is the delay of the wet component in the zenith direction; is the dry component delay in the zenith direction; is the elevation angle of the satellite relative to the satellite navigation signal simulator; are the correction coefficients for the tropospheric dry and wet components, respectively; and are the absolute temperature and water vapor pressure at standard sea level respectively; is the upper limit of the tropospheric dry component based on sea level; is the simulator height; is the atmospheric pressure; 7) Doppler shift calculation: There is a Doppler shift between the navigation satellite and the satellite navigation signal simulator. In order to restore the real satellite signal, the satellite navigation signal simulator needs to calculate the Doppler shift. According to the existing formula: (13) In the formula, is the satellite's velocity, Is related to the signal transmission frequency The corresponding signal wavelength is is the speed of light, is the signal incident angle, that is, the angle between the receiving end and the signal incident direction; assuming that the signal frequency transmitted by the satellite is , the radial velocity in the direction of connection between the measuring point and the satellite is : （14） in, , are the three-dimensional coordinates of the satellite and the measuring point; , is the three-dimensional velocity of the satellite and the measuring point; The Doppler shift of the signal received by the satellite navigation signal simulator at the measurement point and the navigation satellite signal frequency are: （15） （16） It can be seen from this that the Doppler shift is in the opposite direction to the relative motion, and the actual frequency is the frequency value after subtracting the Doppler shift; 8) Pseudorange calculation; Pseudorange can be divided into two parts: the geometric distance between the satellite and the satellite navigation signal simulator; various error values ​​superimposed on the true distance; the pseudorange calculation formula is: （17） in: It is a simulator To Satellite The pseudorange generated value of It is a simulator To Satellite The geometric distance between is the signal reception moment; is the time when the signal is transmitted; is the satellite clock error; is the simulator clock error; is the ionospheric delay; It is the tropospheric delay; is the pseudorange measurement error caused by the multipath effect; is the relativistic effect error; It is the error caused by the satellite phase center deviating from the star's center of mass; is the observation error; Satellite navigation signal simulator at sampling time When receiving satellite signals, the satellite position is ; Since it takes a while for the signal to propagate in space , the signal is The satellite position at the time of transmission is , that is, the signal is from The distance corresponding to the measured time of signal transmission from this position to the simulator is the pseudorange. : （18） in, is the speed of light, which can be taken as ; Since satellites are high-speed moving objects, their positions and There is a big difference. According to formula (17), the calculation of pseudorange yes Navigation satellites at all times The geometric distance between the satellite navigation signal simulators at the moment, when the satellite navigation signal simulator is positioned, is the known signal reception time, and the signal transmission time is unknown and can only be determined by Calculated, and It also needs to be calculated according to formula (18). It can be seen that the calculation equation (17) of the pseudorange is a transcendental equation and needs to be solved iteratively. At the same time, due to the pseudorange Including multiple error terms, the accuracy of the error terms is limited, which leads to errors in the calculation of pseudorange. Therefore, the solution of pseudorange needs to go through multiple iterations to get accurate , and then obtain the accurate pseudorange ; 9) Carrier phase delay calculation; The carrier phase value corresponding to a certain moment cannot be directly measured. The carrier phase observation value is the phase difference between the reference carrier signal generated by the satellite navigation signal simulator at a certain moment and the satellite carrier signal received at this time; Since the carrier phase and the pseudo code are on the same signal, the error term introduced by the signal during propagation is the same for the carrier phase and the pseudorange. Therefore, in the carrier phase calculation, the error term that needs to be considered is the same as that in the pseudorange calculation. The formula is: （19） in: It is a simulator To Satellite The carrier phase delay of It is a simulator To Satellite The geometric distance between is the signal reception moment; is the time when the signal is transmitted; is the satellite clock error; is the simulator clock error; is the ionospheric delay; It is the tropospheric delay; is the pseudorange measurement error caused by the multipath effect; is the relativistic effect error; It is the error caused by the satellite phase center deviating from the star's center of mass; is the observation error; is the carrier wavelength, is the number of integral cycles of the carrier phase, It is the less than full cycle portion of the carrier phase; 10) Calculate the frequency control word; when running in real time, FPGA sends an interrupt to DSP every 8ms. After DSP enters the interrupt, it calculates the frequency control word of each channel and sends it to FPGA, which accumulates it. 11) Monitor and correct FPGA accumulation results; when FPGA sends an interrupt, DSP accurately calculates the current simulation time, obtains the navigation message count, ms count, code chip count and code phase, compares the corresponding counts and code phases of FPGA and DSP, and if the error exceeds the threshold, fine-tunes the frequency control word to correct the error until the error is below the threshold; 12) The GNSS signal simulation of the satellite navigation system is completed.

6. The control method according to claim 5, characterized in that: By calculating the center of gravity of the flight simulation platform, it meets the 15% —29% The aircraft center of gravity limit requirements, so as to simulate the center of gravity position of the real aircraft, the steps are: 1) Determination of the coordinate system of the flight simulation platform: Select The coordinate system is the coordinate system of the platform, where It is the origin of the coordinate system, which is located 1 meter backward from the nose along the horizontal line of the aircraft structure; Axis through origin Along the aircraft construction horizontal line, backward is positive; Axis through origin , in the plane of symmetry of the aircraft, and The axis is vertical, with upward being positive; 2) Flight simulation platform weighing and center of gravity position calculation: Push the flight simulation platform onto the weighbridge, and make the front and main landing gear wheels stop at the center of each weighbridge. The placement of the weighbridge should be consistent with the position of the landing gear wheels. Record the readings of each scale and you will get: （20） in, is the platform weight; The reading of the nose landing gear wheel scale when it is in horizontal state; It is the sum of the readings of the main landing gear wheel scales when in horizontal state; is the distance between the front landing gear wheel and the origin of the coordinate system Towards distance; is the distance between the main landing gear wheels and the origin of the coordinate system Towards distance; is the center of gravity of the platform and the origin of the coordinate system Towards distance; is the center of gravity of the platform and the origin of the coordinate system Distance, since there is no major change in the flight simulation platform, take the theoretical value ; 3) According to the requirements of the remote sensing flight mission, simulate the state of increasing or decreasing the remote sensing equipment, onboard facilities, fuel, and crew load on the remote sensing flight simulation platform in the ground laboratory, and calculate the center of gravity position of the flight simulation platform under this state: （21） in, is the center of gravity of the platform and the origin of the coordinate system Towards distance; It is the weight of the added remote sensing equipment, onboard facilities, fuel, and passenger load; The added remote sensing equipment, onboard facilities, fuel, and crew loads are respectively related to the origin of the coordinate system. Towards distance; It is the weight of reduced remote sensing equipment, onboard facilities, fuel, and passenger load; The reduced loads of remote sensing equipment, onboard facilities, fuel, and crew are respectively related to the origin of the coordinate system. Towards distance; It is the sum of the weight of the added remote sensing equipment, onboard facilities, fuel, and passenger load; It is the sum of the reduced weight of remote sensing equipment, onboard facilities, fuel, and passenger load. All can be measured; is the center of gravity of the platform and the origin of the coordinate system Distance, since there is no major change in the flight simulation platform, take the theoretical value ; 4) Calculate the position of the aircraft's center of gravity on the mean aerodynamic chord: （22） in, is the distance between the leading edge of the mean aerodynamic chord and the origin of the coordinate system Towards distance; is the distance between the leading edge of the mean aerodynamic chord and the origin of the coordinate system Towards distance; is the mean aerodynamic chord and The angle of the axes; It is the position of the projection of the aircraft's center of gravity on the mean aerodynamic chord to the leading edge of the mean aerodynamic chord; It is the vertical distance from the center of gravity of the aircraft to the mean aerodynamic chord; 5) Mean aerodynamic chord ( ) conversion; the center of gravity of the flight simulation platform needs to be converted into the mean aerodynamic chord percentage: （23） in, , is the mean aerodynamic chord percentage of the flight platform's center of gravity, is the mean aerodynamic chord, which is a known value; 6) Check the calculation results of the center of gravity. If The results meet 15% —29% If the aircraft center of gravity limit requirement is met, the flight simulation platform successfully simulates the center of gravity state of the real aircraft. Otherwise, it is necessary to check and correct the combination of loads on the platform and recalculate the center of gravity position of the flight simulation platform.

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