An attitude calculation method and device for an in-atmosphere aircraft

Abstract

The present invention provides a method and device for calculating the attitude of an in-atmosphere aircraft, relating to the technical field of integrated navigation technology for space aircraft. The method includes: taking the moment when the GNSS initial valid measurement is obtained as the reference moment, and using a buffer matrix to store the aircraft parameters in the frozen inertial system and the measurement parameters of the GNSS integrated navigation at the reference moment; within the update period of each INS inertial navigation system, updating the aircraft parameters in the frozen inertial system at the current moment, and within the update period of the GNSS integrated navigation, storing the aircraft parameters in the frozen inertial system updated in the current period and the measurement parameters of the GNSS integrated navigation in the current period; based on the stored aircraft parameters in the frozen inertial system and the measurement parameters of the GNSS integrated navigation, updating the data in the sliding window, so as to recursively calculate the attitude matrix of the aircraft at the current moment by using the updated sliding window data. This solution can quickly and accurately perform real-time calculation of the attitude of the aircraft.

Description

Technical Field

[0001] The present invention relates to the technical field of integrated navigation of space vehicles, and particularly to an attitude calculation method and device for an in-atmosphere vehicle. Background Art

[0002] For the navigation system of an in-atmosphere vehicle, generally, a global navigation satellite system (GNSS) and an inertial navigation device (INS) are required for integrated navigation. However, in the traditional aviation navigation field, when using GNSS and INS integrated navigation to correct attitude navigation, a high-dimensional observation equation needs to be established for filtering and solution, and the high-dimensional navigation filtering solution consumes a huge amount of computing power.

[0003] In related technologies, the navigation systems in the aviation field often configure a separate computer for only navigation solution, so they can bear the corresponding complexity. However, for an orbital reentry vehicle, due to the high cost of entering space and the high demand for reliability, it is required that the navigation, guidance, and control algorithms be integratedly designed and integrated in the same computer. Moreover, for a vehicle flying at high speed into the atmosphere, a very high control bandwidth is usually required, and the high-bandwidth control requirement requires the control period of the control system computer to be as short as possible. Therefore, the control system expects the computing complexity of the navigation solution to be as small as possible.

[0004] Therefore, there is an urgent need to provide an attitude calculation method and device for an in-atmosphere vehicle. Summary of the Invention

[0005] To solve the problem of large attitude calculation amount in traditional GNSS / INS integrated navigation, the present invention provides an attitude calculation method and device for an in-atmosphere vehicle.

[0006] In a first aspect, the present invention provides an attitude calculation method for an in-atmosphere vehicle, and the method includes:

[0007] Taking the moment of the initial effective measurement of GNSS in GNSS integrated navigation as the reference moment, and using a cache matrix to store the vehicle parameters and the measurement parameters of GNSS integrated navigation in the frozen inertial system at the reference moment; wherein, the vehicle parameters include the attitude quaternion and the velocity increment of the vehicle.

[0008] Within each update period of the INS inertial navigation system, updating the vehicle parameters of the frozen inertial system at the current moment, and within the update period of GNSS integrated navigation, storing the updated vehicle parameters of the frozen inertial system in the current period and the measurement parameters of GNSS integrated navigation in the current period; wherein, the measurement parameters include the velocity of the vehicle in the frozen inertial system and the effective time of GNSS integrated navigation.

[0009] Based on the aircraft parameters of the stored frozen inertial system and the measurement parameters of GNSS integrated navigation, update the data in the sliding window, so as to recursively calculate the attitude matrix of the aircraft at the current moment by using the updated sliding window data.

[0010] In a second aspect, the present invention further provides an attitude calculation device for an aircraft within the atmosphere. The device includes:

[0011] A storage unit, configured to use the moment when the GNSS initial valid measurement in the GNSS integrated navigation is taken as the reference moment, and store the aircraft parameters of the initialized frozen inertial system and the measurement parameters of the GNSS integrated navigation at the reference moment by using a cache matrix; wherein, the aircraft parameters include the attitude quaternion and velocity increment of the aircraft.

[0012] An update unit, configured to update the aircraft parameters of the frozen inertial system at the current moment within each update period of the INS inertial navigation system, and store the aircraft parameters of the updated frozen inertial system and the measurement parameters of the GNSS integrated navigation in the current period within the update period of the GNSS integrated navigation; wherein, the measurement parameters include the velocity of the aircraft in the frozen inertial system and the valid time of the GNSS integrated navigation.

[0013] A calculation unit, configured to update the data in the sliding window based on the stored aircraft parameters of the frozen inertial system and the measurement parameters of the GNSS integrated navigation, so as to recursively calculate the attitude matrix of the aircraft at the current moment by using the updated sliding window data.

[0014] In a third aspect, the present invention further provides a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, the method described in any embodiment of this specification is implemented.

[0015] In a fourth aspect, the present invention further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed on a computer, the computer is made to execute the method described in any embodiment of this specification.

[0016] On the other hand, the present application further provides a computer program product. The computer program product includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium, and the processor executes the computer program, so that the computer device executes the method described in any one of the above first aspects.

[0017] The present invention provides a method for calculating the attitude of an in - atmosphere aircraft. By introducing the assumption of frozen inertia, taking the moment of the initial effective measurement of GNSS in GNSS integrated navigation as the reference moment, continuously updating and storing the aircraft parameters in the frozen inertial system and the measurement parameters of integrated navigation. Then, using the acceleration integration information of the INS inertial navigation system within a period of time and the measurement parameters of GNSS integrated navigation, a method for calculating the attitude matrix based on the instantaneous inertial system is constructed, and through continuous iteration with a sliding window, the real - time attitude information at any moment is obtained, and the inertial navigation attitude is corrected, thereby achieving fast and accurate calculation of the aircraft attitude at the current moment. Brief Description of the Drawings

[0018] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0019] Figure 1 is a flowchart of a method for calculating the attitude of an in - atmosphere aircraft provided by an embodiment of the present invention;

[0020] Figure 2 is a graph showing the changes in height and speed corresponding to the cross - section of the aircraft during flight provided by an embodiment of the present invention; wherein, the ordinate in the upper graph is height, with the unit of m, and the ordinate in the lower graph is speed, with the unit of m / s;

[0021] Figure 3 is a graph showing the changes in Euler angles corresponding to the cross - section of the aircraft during flight provided by an embodiment of the present invention; wherein, pitch represents the pitch angle, yaw represents the yaw angle, and roll represents the roll angle;

[0022] Figure 4 is a schematic diagram of the error results obtained by calculating the pure inertial navigation attitude provided by an embodiment of the present invention; wherein, Δpitch represents the pitch angle error, Δyaw represents the yaw angle error, and Δroll represents the roll angle error;

[0023] Figure 5 is a schematic diagram of the attitude navigation error results obtained by using the method in the embodiment of the present invention provided by an embodiment of the present invention;

[0024] Figure 6 is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0025] Figure 7 is a structural diagram of an attitude calculation device for an in - atmosphere aircraft provided by an embodiment of the present invention. Detailed implementation manners

[0026] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Apparently, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without making creative efforts shall fall within the protection scope of the present invention.

[0027] The following describes the specific implementation manners of the above concept.

[0028] Please refer to Figure 1 , an attitude calculation method for an in-atmosphere aircraft is provided in an embodiment of the present invention. The method includes:

[0029] Step 100: Use the moment when the GNSS initial effective measurement in the GNSS integrated navigation is taken as the reference moment, and store the aircraft parameters in the frozen inertial system and the measurement parameters of the GNSS integrated navigation at the reference moment by using a cache matrix; wherein, the aircraft parameters include the attitude quaternion and the velocity increment of the aircraft.

[0030] Step 102: Update the aircraft parameters in the frozen inertial system at the current moment within each update period of the INS inertial navigation system, and store the updated aircraft parameters in the frozen inertial system and the measurement parameters of the GNSS integrated navigation in the current period within the update period of the GNSS integrated navigation; wherein, the measurement parameters include the velocity of the aircraft in the frozen inertial system and the effective time of the GNSS integrated navigation.

[0031] Step 104: Update the data in the sliding window based on the stored aircraft parameters in the frozen inertial system and the measurement parameters of the GNSS integrated navigation, so as to recursively calculate the attitude matrix of the aircraft at the current moment by using the updated sliding window data.

[0032] In the embodiment of the present invention, by introducing the frozen inertia hypothesis, with the moment when the GNSS initial effective measurement in the GNSS integrated navigation is taken as the reference moment, and continuously updating and storing the aircraft parameters in the frozen inertial system and the measurement parameters of the integrated navigation, then, by using the acceleration integration information of the INS inertial navigation system within a period of time and the measurement parameters of the GNSS integrated navigation, an attitude matrix calculation method based on the instantaneous inertial system is constructed, and through continuous iteration of the sliding window, real-time attitude information at any moment is obtained, and the inertial navigation attitude is corrected, thereby realizing the fast and accurate calculation of the aircraft attitude at the current moment.

[0033] The basic principle of the embodiment of the present invention draws on the method proposed by Professor Qin Yongyuan et al. to solve the initial alignment under the condition of a shaking base. This method applies the inertial freezing hypothesis and determines the attitude of the carrier with the aid of gravity information. In this embodiment, the freezing inertial hypothesis is extended to a carrier vehicle (aircraft) flying within the atmosphere. Considering that general aircraft are equipped with a basic configuration of GNSS / INS integrated navigation, it can be considered that when GNSS is effective, the position and velocity information of the carrier are known information. Therefore, without establishing and solving a high-dimensional coupled filtering matrix, the real-time inertial attitude of the carrier can be quickly calculated.

[0034] The derivation of the basic principle in the embodiment of the present invention is as follows:

[0035] In a moving environment, given the position of the carrier (i.e., the aircraft) at any moment in the navigation inertial system, it is necessary to determine the attitude matrix of the carrier relative to the navigation inertial system.

[0036] The attitude matrix of the carrier relative to the navigation inertial system can be written as the product of the following two matrices:

[0037]

[0038] In the formula, is the transformation matrix between the freezing inertial system (denoted as ) of the aircraft at the initial moment and the aircraft coordinate system (denoted as b). At the initial moment, is the identity matrix. After the start of navigation, the attitude quaternion is updated and solved using the angular velocity information output by the gyro in real time. is the attitude matrix of the freezing inertial system of the aircraft relative to the navigation inertial system (denoted as i) at the initial moment.

[0039] Furthermore, the velocity of the aircraft in the freezing inertial system can be expressed in the following integral form:

[0040]

[0041] In the formula, , are respectively the velocity vectors of the aircraft in the freezing inertial system at time , is the gravity vector in the inertial system,

[0042] When the GNSS integrated navigation system signal is valid, the aircraft can obtain , the velocity vector of the freezing inertial system at time ,

[0043] Let , let ,

[0044] Then the left side of the above formula can be rewritten as:

[0045]

[0046] Let , ,

[0047] Respectively select , moments ( ), and respectively obtain:

[0048] ;

[0049] ;

[0050] By performing double-vector attitude determination through the above two formulas, the transformation matrix of the solidified inertial system of the aircraft at the initial moment relative to the inertial system can be determined, and then the attitude matrix of the aircraft at the initial moment relative to the navigation inertial system can be obtained.

[0051] Regarding step 100:

[0052] In the embodiment of the present invention, when the initial measurement of GNSS in the GNSS integrated navigation is valid, the first-step calculation is started, and the current time t is assigned to the initial reference time :

[0053]

[0054] Initialize the attitude quaternion of the aircraft in the solidified inertial system: , where is the attitude quaternion of the aircraft in the solidified inertial system and is a 4D array;

[0055] Initialize the velocity increment accumulated by the accelerometer and the velocity increment accumulated by gravity , are both 3D vectors.

[0056] After that, a cache data matrix is established to store the aircraft parameters in the solidified inertial system and the measurement parameters of the GNSS integrated navigation at the reference moment, specifically as follows:

[0057] Store the attitude quaternion of the solidified inertial system: ; where, is an N 4D matrix, and N is the defined sliding window length;

[0058] Store the accumulated speed increment of the accelerometer: is an N 3D matrix;

[0059] Store the accumulated speed increment of gravity: is an N 3D matrix;

[0060] Store the speed of the aircraft in the inertial system measured by the GNSS integrated navigation at the reference time: is an N 3D matrix, is the speed vector of the aircraft in the inertial system measured by the GNSS integrated navigation at the reference time, and this variable is input by the external navigation;

[0061] Store the effective time of the GNSS initial measurement in the initial GNSS integrated navigation: ;

[0062] And record the storage amount of the current data .

[0063] Regarding step 102:

[0064] In some embodiments, step 102 includes:

[0065] Update the attitude quaternion of the aircraft in the current cycle according to the attitude quaternion of the previous cycle and the angular rate of the aircraft measured by the INS inertial navigation system in the current cycle;

[0066] Update the accumulated speed increment of the accelerometer of the aircraft in the current cycle according to the accumulated speed increment of the accelerometer of the previous cycle and the acceleration of the aircraft measured by the INS inertial navigation system in the current cycle;

[0067] Update the accumulated speed increment of gravity of the aircraft in the current cycle according to the accumulated speed increment of gravity of the previous cycle and the position of the aircraft measured by the GNSS integrated navigation in the geocentric inertial system in the current cycle.

[0068] The specific process of updating the aircraft parameters in the fixed inertial system is as follows:

[0069] The angular rate of the aircraft measured by the INS inertial navigation system in each cycle is and the acceleration of the aircraft is , and the update period of the INS inertial navigation system is .

[0070] First, calculate the angle increment of the aircraft according to the following formula :

[0071]

[0072] After that, update the attitude quaternion of the current cycle according to the following formula :

[0073]

[0074]

[0075]

[0076] Among them, is the attitude quaternion of the aircraft in the previous cycle;

[0077] After that, first calculate the transformation matrix between the fixed inertial system of the aircraft and the aircraft coordinate system in the current cycle according to the following formula :

[0078] On this basis, update the velocity increment of the accelerometer integration in the current cycle according to the following formula :

[0079]

[0080] Among them, is the velocity increment of the accelerometer integration in the previous cycle, is a four-dimensional array, , , and represent the numbers of each dimension respectively;

[0081] Finally, update the velocity increment of the gravity integration in the current cycle according to the following formula :

[0082]

[0083] Among them, is the velocity increment of the gravity integration in the previous cycle, is the gravitational constant, is the position of the aircraft in the geocentric inertial system measured by the GNSS integrated navigation, and this variable is input by the GNSS integrated navigation, , R is the distance between the aircraft and the earth's center, and the update period of the INS inertial navigation system is .

[0084] In the embodiments of the present invention, by means of pure inertial recursion, the continuous update of the aircraft parameters in the fixed inertial system can be achieved.

[0085] Regarding step 104:

[0086] In some embodiments, in step 104, updating the data in the sliding window based on the stored aircraft parameters in the fixed inertial system and the measurement parameters of the GNSS navigation system includes:

[0087] If the amount of data stored in the current cache matrix is greater than the first threshold and less than the length of the sliding window, calculate the attitude of the aircraft based on the data stored in the current cache matrix;

[0088] If the amount of data stored in the current cache matrix is the same as the length of the sliding window, use the sliding window to update the data stored in the cache matrix in sequence according to the time series, and calculate the attitude of the aircraft in sequence according to the data stored in the updated cache matrix at the current moment.

[0089] Since the navigation period of the aircraft is generally less than the update frequency of the GNSS navigation system, and the GNSS also has inherent characteristics such as discontinuous signals, there are a large number of navigation time periods in the invalid state of the GNSS navigation system. However, only by correcting the position and speed of the aircraft during the valid period of the GNSS navigation system can the further effective calculation of the attitude information of the aircraft be achieved. Therefore, in the embodiments of the present invention, during the update period of the GNSS integrated navigation, the aircraft parameters in the fixed inertial system updated in the current period and the measurement parameters of the GNSS integrated navigation in the current period are stored, specifically as follows:

[0090] Store the attitude quaternion of the fixed inertial system in the current period: ;

[0091] Store the speed increment accumulated by the accelerometer in the current period: ;

[0092] Store the speed increment accumulated by gravity in the current period: ;

[0093] Store the speed in the inertial system determined by the GNSS integrated navigation at the corresponding moment in the current period: ;

[0094] Store the effective time of the initial measurement of the GNSS in the GNSS integrated navigation in the current period: ;

[0095] And record the current data storage amount , until the amount of data in the cache matrix is the same as the preset window length N.

[0096] In some specific embodiments, when the amount of data in the cache matrix is the same as the length of the sliding window, the sliding window specifically updates the data in the cache matrix in the following manner:

[0097] , i = 0 to , all the currently stored data are shifted forward by one position;

[0098] Store the quaternion of the inertial frame of the current cycle's coagulation: ;

[0099] , i = 0 to , all the currently stored data are shifted forward by one position;

[0100] Store the velocity increment of the current cycle's cumulative acceleration: ;

[0101] , i = 0 to , all the currently stored data are shifted forward by one position;

[0102] Store the velocity increment of the current cycle's cumulative gravity: ;

[0103] , i = 0 to , all the currently stored data are shifted forward by one position;

[0104] Store the velocity in the inertial frame measured by the GNSS integrated navigation at the current moment: ;

[0105] , i = 0 to , all the currently stored data are shifted forward by one position;

[0106] Store the effective time of the GNSS initial measurement in the GNSS integrated navigation of the current cycle: ;

[0107] And record the current data storage amount .

[0108] In the embodiments of the present invention, if the amount of data stored in the current cache matrix is small, specifically greater than the first threshold (for example, it can be 2) and less than the length N of the sliding window, the attitude of the aircraft is calculated based on the data stored in the current cache matrix. As time progresses, the error accumulated by pure inertial navigation will gradually increase, and if the time is too short, the small difference between the two selected vectors will also affect the accuracy. Therefore, in the embodiments of the present invention, in the case of a relatively long navigation duration during the movement of the aircraft, the sliding window is used to sequentially slide and receive new stored data in time series, discard the oldest stored data, and successively obtain the absolute attitude calculation result based on the inertial navigation integration information of a period of time according to the inertial navigation integration information of a segment of data within the sliding window at the current moment and the real-time position and velocity information given by the GNSS integrated navigation.

[0109] In some embodiments, the attitude of the aircraft at the current moment is calculated as follows:

[0110] Select an intermediate point moment in the current cache matrix, and based on the data stored in the current cache matrix, respectively derive the incremental vectors generated by the incremental variables of the accelerometer integrated from the reference moment to the current moment and the intermediate point moment in the fixed inertial system;

[0111] According to the incremental vectors generated by the incremental variables of the accelerometer at the current moment and the intermediate point moment in the fixed inertial system, obtain the attitude matrix of the fixed inertial system at the reference moment relative to the navigation inertial system;

[0112] According to the transformation matrix between the fixed inertial system at the current moment and the aircraft inertial system, the attitude matrix of the fixed inertial system at the reference moment relative to the navigation inertial system, and the conversion matrix of the fixed inertial system at the reference moment relative to the aircraft inertial system, obtain the attitude matrix of the aircraft relative to the navigation inertial system at the current moment.

[0113] In the embodiments of the present invention, when calculating the attitude of the aircraft, first based on the amount of data stored in the current cache matrix Select an intermediate point moment. If is an even number, the selected intermediate point moment , if is an odd number, the selected intermediate point moment .

[0114] In some specific embodiments, the incremental vectors generated by the incremental variables of the accelerometer at the current moment and the intermediate point moment in the fixed inertial system are determined as follows:

[0115] Respectively calculate the velocity increment of the gravity accumulation from the reference moment to the intermediate point moment and the current moment, and the velocity of the aircraft in the geocentric inertial system;

[0116] Calculate the velocity increment vectors generated by factors other than gravity accumulated from the reference time to the intermediate point time and the current time, respectively, based on the velocity increment accumulated by gravity up to the intermediate point time and the current time, and the velocity of the aircraft in the geocentric inertial system;

[0117] Calculate the attitude matrix of the frozen inertial system at the reference time relative to the navigation inertial system based on the velocity increment generated by factors other than the gravity and the attitude quaternion of the aircraft;

[0118] Calculate the increment vectors generated by the accelerometer cumulative variables accumulated from the reference time to the intermediate point time and the current time in the frozen inertial system, respectively, based on the attitude matrix of the frozen inertial system at the reference time relative to the navigation inertial system and the stored accelerometer cumulative velocity increment.

[0119] In the embodiments of the present invention, when performing attitude calculation, the velocity increments generated by gravity accumulated from the reference time to the intermediate point time and the current time are calculated based on the following formula 、 :

[0120]

[0121]

[0122] wherein, are all three-dimensional vectors, is the velocity increment accumulated by gravity stored at the intermediate point time in the cache matrix, is the velocity increment accumulated by gravity stored at the reference time in the cache matrix, is the velocity increment accumulated by gravity stored at the current time in the cache matrix;

[0123] Calculate the inertial system velocities obtained by GNSS integrated navigation from the reference time to the intermediate point time and the current time through the following formula 、 :

[0124]

[0125]

[0126] wherein, are all three-dimensional vectors, is the velocity of the aircraft measured by GNSS integrated navigation stored at the intermediate point time in the cache matrix in the inertial system, is the velocity of the aircraft measured by GNSS integrated navigation stored at the reference time in the cache matrix in the inertial system, is the velocity of the aircraft measured by GNSS integrated navigation stored at the current time in the cache matrix in the inertial system;

[0127] Thus, the velocity increments generated by forces other than gravity are calculated up to the intermediate point time and the current time. , :

[0128]

[0129]

[0130] Wherein, , are all three-dimensional vectors;

[0131] The above two velocity increments are normalized in the following manner:

[0132]

[0133]

[0134]

[0135]

[0136]

[0137] Let the first vector matrix (3×3 dimension) ;

[0138] Where , and represent the numbers of each dimension in the three-dimensional vector respectively;

[0139] When calculating the increment vector , ( , are all three-dimensional vectors) generated by the accelerometer cumulative variable from the reference time to the intermediate point time and the current time in the solidification inertial system, it is first necessary to calculate the attitude matrix of the solidification inertial system relative to the navigation inertial system through the following formula:

[0140] Let , then:

[0141]

[0142] Thus, based on the calculation of , can be achieved through the following formula:

[0143]

[0144]

[0145] Wherein, is the increment of the accumulated acceleration stored in the cache matrix at the intermediate point in time, is the increment of the accumulated acceleration stored in the cache matrix at the reference time, is the increment of the accumulated acceleration stored in the cache matrix at the current time;

[0146] Furthermore, the above two increments of acceleration are normalized in the following manner:

[0147]

[0148]

[0149]

[0150]

[0151]

[0152] Let the second vector matrix (3×3 dimension)

[0153] Therefore, the attitude matrix of the solidified inertial system defined at the reference time relative to the navigation inertial system can be obtained according to the first vector matrix and the second vector matrix through the following formula :

[0154]

[0155] Finally, according to the attitude matrix of the solidified inertial system at the reference time relative to the navigation inertial system , the attitude matrix of the solidified inertial system at the reference time relative to the navigation inertial system and the transformation matrix between the solidified inertial system and the vehicle inertial system at the current time , the attitude matrix of the vehicle relative to the navigation inertial system at the current time can be further solved through the following formula :

[0156] .

[0157] In summary, in the embodiments of the present invention, taking an initial moment as a reference, using the velocity increment vector generated by pure inertial navigation recursion to the current moment, and combining the velocity increment vector at an intermediate moment, through double-vector attitude determination, the attitude of the inertial system at the reference moment is first calculated. Combining the relative attitude increment obtained by inertial navigation during this period, the attitude of the current moment relative to the inertial system can be obtained. At the same time, aiming at the characteristic that the error accumulated by pure inertial navigation will gradually increase, in the embodiments of the present invention, by establishing and maintaining a sliding window data, the duration of the inertial navigation recursion data used is controlled, thus avoiding the problem that the accuracy diverges continuously with the cumulative duration of inertial navigation.

[0158] Moreover, the method based on double-vector attitude determination in the embodiments of the present invention is simple, has low requirements for the performance of the aircraft computer, can realize accurate real-time calculation of the aircraft attitude, and this method can be applied to the attitude calculation of aircraft within the atmosphere with high-reliability integrated navigation, guidance, and control calculation requirements.

[0159] It should be noted that in the embodiments of the present invention, when the amount of data stored in the cache matrix is the same as the length of the sliding window, when using the sliding window to update the data in the cache matrix and calculate the attitude of the aircraft, the data at the reference moment in the cache matrix is constantly changing. Therefore, the reference moment actually refers to the first moment in the sliding window.

[0160] In order to verify the effect of the method in the embodiments of the present invention, in the embodiments of the present invention, for an aircraft that glides and re-enters the atmosphere starting from an altitude of 120 km at a speed of 7800 m / s, the total flight duration is 2000 seconds.

[0161] The initial three-axis attitude errors are 0.1 degrees respectively; the gyro zero bias is considered 0.1 degree / hour; the gyro zero bias stability is 0.05 degree / 10 hours; the accelerometer zero bias is 5×e -5 g 0 (g 0 is the magnitude of the standard gravitational acceleration);

[0162] After the altitude is lower than 50 km, stable GNSS measurement information is obtained.

[0163] Compare the attitude calculated by the method in the embodiments of the present invention with the pure inertial navigation attitude measurement result. The comparison result is as Figures 2 to 5 shown. It can be seen from the figure that compared with the state where the attitude navigation error diverges continuously with time under the pure inertial navigation condition, after using the method in the embodiments of the present invention for attitude correction, the Euler angle attitude error converges rapidly, which shows that the method in the embodiments of the present invention can realize fast calculation of the aircraft attitude.

[0164] As Figure 6 、 Figure 7As shown in the figure, an embodiment of the present invention provides an attitude calculation device for an in-atmosphere aircraft. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. At the hardware level, as Figure 6 shown, it is a hardware architecture diagram of a computing device where the attitude calculation device for an in-atmosphere aircraft provided by an embodiment of the present invention is located. In addition to Figure 6 the processor, memory, network interface, and non-volatile memory shown, the computing device where the device is located in the embodiment usually may also include other hardware, such as a forwarding chip responsible for processing packets, etc. Taking software implementation as an example, as Figure 7 shown, as a logically meaningful device, it is formed by the CPU of its computing device reading the corresponding computer program in the non-volatile memory into the memory for operation.

[0165] An attitude calculation device for an in-atmosphere aircraft provided in this embodiment, the device includes:

[0166] A storage unit 701, configured to use the moment of the initial effective measurement of GNSS in GNSS integrated navigation as the reference moment, and store the aircraft parameters in the frozen inertial system and the measurement parameters of GNSS integrated navigation at the reference moment by using a cache matrix; wherein, the aircraft parameters include the attitude quaternion and velocity increment of the aircraft;

[0167] An update unit 702, configured to update the aircraft parameters in the frozen inertial system at the current moment within each update period of the INS inertial navigation system, and store the updated aircraft parameters in the frozen inertial system and the measurement parameters of GNSS integrated navigation in the current period within the update period of GNSS integrated navigation; wherein, the measurement parameters include the velocity of the aircraft in the frozen inertial system and the valid time of GNSS integrated navigation;

[0168] A calculation unit 703, configured to update the data in the sliding window based on the stored aircraft parameters in the frozen inertial system and the measurement parameters of GNSS integrated navigation, so as to recursively calculate the attitude matrix of the aircraft at the current moment by using the updated sliding window data.

[0169] In an embodiment of the present invention, the storage unit 701 may be configured to execute step 100 in the above method embodiment, the update unit 702 may be configured to execute step 102 in the above method embodiment, and the calculation unit 703 may be configured to execute step 104 in the above method embodiment.

[0170] In an embodiment of the present invention, the velocity increment of the aircraft includes the velocity increment accumulated by the accelerometer and the velocity increment accumulated by gravity;

[0171] During each update cycle of each INS inertial navigation system, the update unit 702 updates the aircraft parameters of the frozen inertial system in the following manner:

[0172] Update the attitude quaternion of the aircraft in the current cycle according to the attitude quaternion of the previous cycle and the angular rate of the aircraft measured by the INS inertial navigation system in the current cycle;

[0173] Update the velocity increment accumulated by the accelerometer in the current cycle according to the velocity increment accumulated by the accelerometer in the previous cycle and the acceleration of the aircraft measured by the INS inertial navigation system in the current cycle;

[0174] Update the velocity increment accumulated by gravity in the current cycle according to the velocity increment accumulated by gravity in the previous cycle and the position of the aircraft in the geocentric inertial system measured by the GNSS integrated navigation in the current cycle.

[0175] In an embodiment of the present invention, when the update unit 702 updates the data in the sliding window based on the stored aircraft parameters of the frozen inertial system and the measurement parameters of the GNSS navigation system, it is used to perform the following operations:

[0176] If the amount of data stored in the current cache matrix is greater than the first threshold and less than the length of the sliding window, calculate the attitude of the aircraft based on the data stored in the current cache matrix;

[0177] If the amount of data stored in the current cache matrix is the same as the length of the sliding window, use the sliding window to update the data stored in the cache matrix in sequence according to the time series, and calculate the attitude of the aircraft according to the data stored in the updated cache matrix at the current moment in sequence.

[0178] In an embodiment of the present invention, in the calculation unit 703, the attitude of the aircraft at the current moment is calculated in the following manner:

[0179] Select an intermediate point moment in the current cache matrix, and based on the data stored in the current cache matrix, recursively calculate the increment vectors generated by the accelerometer cumulative variables at the current moment and the intermediate point moment in the frozen inertial system from the reference moment;

[0180] Obtain the attitude matrix of the frozen inertial system relative to the navigation inertial system at the reference moment according to the increment vectors generated by the accelerometer cumulative variables at the current moment and the intermediate point moment in the frozen inertial system;

[0181] Obtain the attitude matrix of the aircraft relative to the navigation inertial system at the current moment according to the transformation matrix between the frozen inertial system and the aircraft inertial system at the current moment, the attitude matrix of the frozen inertial system relative to the navigation inertial system at the reference moment, and the conversion matrix of the frozen inertial system relative to the aircraft inertial system at the reference moment.

[0182] In an embodiment of the present invention, in the calculation unit 703, the incremental vector generated by the cumulative variables of the current moment and the intermediate point moment in the solidification inertial system is determined in the following manner:

[0183] Calculate the velocity increments of the gravity accumulation from the reference moment to the intermediate point moment and the current moment respectively, and the velocity of the aircraft in the geocentric inertial system;

[0184] According to the velocity increments of the gravity accumulation from the reference moment to the intermediate point moment and the current moment, and the velocity of the aircraft in the geocentric inertial system, calculate the incremental velocity vectors generated by factors other than gravity from the reference moment to the intermediate point moment and the current moment respectively;

[0185] Based on the incremental velocity generated by factors other than gravity and the attitude quaternion of the aircraft, calculate the attitude matrix of the solidification inertial system at the reference moment relative to the navigation inertial system;

[0186] According to the attitude matrix of the solidification inertial system at the reference moment relative to the navigation inertial system and the stored cumulative velocity increment of the accelerometer, calculate the incremental vectors generated by the cumulative variables of the accelerometer from the reference moment to the intermediate point moment and the current moment in the solidification inertial system respectively.

[0187] In an embodiment of the present invention, in the calculation unit 703, the attitude matrix of the solidification inertial system at the reference moment relative to the navigation inertial system is obtained by multiplying the first vector matrix and the second vector matrix; wherein, the first vector matrix is determined by the incremental velocity generated by factors other than gravity, and the second vector matrix is determined by the incremental vector generated by the cumulative variables of the accelerometer in the solidification inertial system.

[0188] It can be understood that the structure illustrated in the embodiments of the present invention does not constitute a specific limitation on an attitude calculation device for an aircraft within the atmosphere. In other embodiments of the present invention, an attitude calculation device for an aircraft within the atmosphere may include more or fewer components than those shown in the figure, or combine certain components, or split certain components, or have different component arrangements. The components shown in the figure can be implemented in hardware, software, or a combination of software and hardware.

[0189] Regarding the information interaction, execution process, etc. between the modules in the above device, since they are based on the same concept as the method embodiments of the present invention, the specific content can be referred to the description in the method embodiments of the present invention, and will not be elaborated here.

[0190] The embodiments of the present invention also provide a computing device, including a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, it implements an attitude calculation method for an aircraft within the atmosphere in any embodiment of the present invention.

[0191] An embodiment of the present invention also provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the processor is caused to execute a method for calculating the attitude of an in-atmosphere aircraft according to any one of the embodiments of the present invention.

[0192] Specifically, a system or device equipped with a storage medium can be provided, on which software program code for implementing the functions of any one of the above embodiments is stored, and the computer (or CPU or MPU) of the system or device is caused to read and execute the program code stored in the storage medium.

[0193] In this case, the program code read from the storage medium itself can implement the functions of any one of the above embodiments, so the program code and the storage medium storing the program code constitute a part of the present invention.

[0194] Examples of the storage medium for providing the program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Optionally, the program code can be downloaded from a server computer via a communication network.

[0195] In addition, it should be clear that not only can the functions of any one of the above embodiments be realized by executing the program code read by the computer, but also by causing an operating system or the like operating on the computer based on the instructions of the program code to complete part or all of the actual operations.

[0196] In addition, it can be understood that the program code read from the storage medium is written into the memory provided in an expansion board inserted into the computer or into the memory provided in an expansion module connected to the computer, and then based on the instructions of the program code, the CPU or the like installed on the expansion board or the expansion module is caused to execute part and all of the actual operations, so as to realize the functions of any one of the above embodiments.

[0197] An embodiment of the present application also provides a computer-readable storage medium, on which at least one instruction, at least one program, a code set or an instruction set is stored. The at least one instruction, at least one program, a code set or an instruction set is loaded and executed by a processor to implement a method for calculating the attitude of an in-atmosphere aircraft provided by the above method embodiments.

[0198] It should be noted that, in this article, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variant thereof are intended to cover non-exclusive inclusion, such that a process, method, article or device comprising a series of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article or device.

[0199] Those of ordinary skill in the art can understand that all or part of the steps for implementing the above method embodiments can be completed by hardware related to program instructions. The foregoing program can be stored in a computer-readable storage medium. When the program is executed, it performs the steps including the above method embodiments; and the foregoing storage medium includes various media that can store program codes, such as ROM, RAM, magnetic disks or optical discs.

[0200] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them. Although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions described in the foregoing embodiments, or perform equivalent replacements for some of the technical features; and these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A method for calculating the attitude of an aircraft in the atmosphere, characterized in that: include: The time of the initial effective measurement of GNSS in GNSS integrated navigation is used as the reference time, and the buffer matrix is ​​used to store the aircraft parameters of the solidified inertial system and the measurement parameters of GNSS integrated navigation at the reference time; wherein the aircraft parameters include the attitude quaternion and velocity increment of the aircraft; In each update cycle of the INS inertial navigation system, the aircraft parameters of the solidified inertial system at the current moment are updated, and in the update cycle of the GNSS integrated navigation, the aircraft parameters of the solidified inertial system updated in the current cycle and the measurement parameters of the GNSS integrated navigation in the current cycle are stored; wherein the measurement parameters include the speed of the aircraft in the solidified inertial system and the effective time of the GNSS integrated navigation; Based on the stored aircraft parameters of the solidified inertial system and the measurement parameters of the GNSS integrated navigation, the data of the sliding window is updated, so as to recursively calculate the attitude matrix of the aircraft at the current moment using the updated sliding window data.

2. The method according to claim 1, characterized in that The speed increment of the aircraft includes the speed increment accumulated by the additive calculation and the speed increment accumulated by gravity; In each update cycle of the INS inertial navigation system, the aircraft parameters of the solidified inertial system are updated in the following manner: Update the attitude quaternion of the aircraft in the current cycle according to the attitude quaternion of the previous cycle and the angular rate of the aircraft measured by the INS inertial navigation system in the current cycle; Update the accumulated velocity increment of the aircraft in the current cycle according to the accumulated velocity increment of the previous cycle and the aircraft acceleration measured by the INS inertial navigation system in the current cycle; The accumulated velocity increment of the gravity of the aircraft in the current period is updated according to the accumulated velocity increment of the gravity of the previous period and the position of the aircraft in the geocentric inertial system measured by the GNSS integrated navigation in the current period.

3. The method according to claim 1, characterized in that The updating of the data of the sliding window based on the stored aircraft parameters of the solidified inertial system and the measurement parameters of the GNSS navigation system comprises: If the amount of data stored in the current cache matrix is ​​greater than the first threshold and less than the length of the sliding window, the attitude of the aircraft is calculated based on the data stored in the current cache matrix; If the amount of data stored in the current cache matrix is ​​the same as the length of the sliding window, the sliding window is used to update the data stored in the cache matrix in sequence according to the time series, and the attitude of the aircraft is calculated in sequence according to the data stored in the cache matrix updated at the current moment.

4. The method according to claim 3, characterized in that: The current attitude of the aircraft is calculated as follows: Select an intermediate point in the current cache matrix, and based on the data stored in the current cache matrix, recursively calculate the incremental vector generated by the added accumulated variables of the current time and the intermediate point in the solidified inertial system from the reference time; According to the incremental vector generated by the added cumulative variables at the current time and the midpoint time in the solidified inertial system, the attitude matrix of the solidified inertial system relative to the navigation inertial system at the reference time is obtained; According to the transformation matrix between the solidified inertial system and the aircraft inertial system at the current moment, the attitude matrix of the solidified inertial system relative to the navigation inertial system at the reference moment, and the conversion matrix of the solidified inertial system relative to the aircraft inertial system at the reference moment, the attitude matrix of the aircraft relative to the navigation inertial system at the current moment is obtained.

5. The method according to claim 4, characterized in that The incremental vector generated by the added cumulative variables at the current moment and the midpoint moment in the solidification inertial system is determined by the following method: Calculate the velocity increment of the gravity accumulated from the reference time to the midpoint time and the current time, and the velocity of the aircraft in the earth's center inertial system; According to the velocity increment accumulated from the midpoint and the current moment by gravity and the velocity of the aircraft in the Earth's center inertial system, the velocity increment vector generated by gravity accumulated from the reference moment to the midpoint and the current moment is calculated respectively; Based on the velocity increment generated other than the gravity and the attitude quaternion of the aircraft, the attitude matrix of the solidified inertial system relative to the navigation inertial system at the reference time is calculated; According to the attitude matrix of the solidified inertial system relative to the navigation inertial system at the reference time and the stored cumulative velocity increment, the incremental vector generated by the cumulative variables accumulated from the reference time to the midpoint time and the current time in the solidified inertial system is calculated respectively.

6. The method according to claim 5, characterized in that The attitude matrix of the solidified inertial system relative to the navigation inertial system at the reference moment is obtained by multiplying the first vector matrix and the second vector matrix; wherein the first vector matrix is ​​determined by the velocity increment generated other than gravity, and the second vector matrix is ​​determined by the incremental vector generated by the added accumulated variables in the solidified inertial system.

7. An attitude calculation device for an aircraft in the atmosphere, characterized in that: include: A storage unit, used to use the time of initial effective GNSS measurement in GNSS integrated navigation as a reference time, and use a cache matrix to store the aircraft parameters of the initialization solidified inertial system and the measurement parameters of the GNSS integrated navigation at the reference time; wherein the aircraft parameters include the attitude quaternion and velocity increment of the aircraft; An updating unit, used to update the aircraft parameters of the solidified inertial system at the current moment in each update cycle of the INS inertial navigation system, and store the aircraft parameters of the solidified inertial system updated in the current cycle and the measurement parameters of the GNSS integrated navigation in the current cycle in the update cycle of the GNSS integrated navigation; wherein the measurement parameters include the speed of the aircraft in the solidified inertial system and the effective time of the GNSS integrated navigation; The calculation unit is used to update the data of the sliding window based on the stored aircraft parameters of the solidified inertial system and the measurement parameters of the GNSS integrated navigation, so as to recursively calculate the attitude matrix of the aircraft at the current moment using the updated sliding window data.

8. A computing device, comprising a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the method according to any one of claims 1 to 6 is implemented.

9. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to execute the method according to any one of claims 1 to 6.

10. A computer program product, characterized in that The method comprises a computer program, which implements the method according to any one of claims 1 to 6 when being executed by a processor.

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