A method for simulating equivalent signals for gyroscope testing based on Euler discrete model

Abstract

This invention provides a simulation method for test equivalent signals based on the Euler discrete model. It primarily solves the problem of equivalent generation of gyroscope angle signals in test equivalent converters, enabling high-precision reproduction of the complex dynamic angle changes of an aircraft during flight. This overcomes the difficulty of accurately simulating gyroscope angle signals during ground testing. First, using predetermined aircraft pitch channel altitude trajectory data, a fast differentiator is used to solve for the differential data of the altitude and horizontal trajectories, thereby solving for the aircraft pitch angle and pitch velocity data. Then, the Euler method is used to establish a discrete measurement model of the gyroscope, solving for the measured value of the pitch velocity. Considering the influence of error drift, integration is performed to obtain the gyroscope's measurement signal of the pitch angle. This outputs a high-precision result of the gyroscope angle changes during high-altitude flight. This method has the advantages of high simulation accuracy and good agreement with reality.

Description

Technical Field

[0001] This invention relates to the field of test equivalents, and more specifically, to a method for simulating the equivalent signal of a gyroscope test based on the Euler discrete model. Background Technology

[0002] Large and complex systems are expensive. To ensure their normal operation, testing and technical preparation before use are essential for the safe completion of missions. For example, large rockets, missiles, and aircraft require multiple tests and launch preparations before launch. Therefore, testing systems for large and complex systems are also very complex and expensive. To verify the operation of the testing system or train technicians to operate it, a test equivalent is needed to replace the real complex system, such as a rocket, missile, or aircraft, to support the testing system. Thus, test equivalents are an emerging research branch in the field of high-precision complex systems. Currently, the mainstream research approach for gyroscope test equivalents is to use software to simulate the gyroscope's on / off signals and power-on conditions. The drawback of this method is that it can only roughly simulate the physical on / off faults of the gyroscope and cannot simulate the real relationship between the aircraft's attitude angle and trajectory. Of course, physical gyroscope systems can also be used, but this approach has the disadvantages of being expensive and easily damaged after repeated operation and power-on tests. Based on the above background reasons, this invention proposes a method that uses an Euler discrete model to establish a realistic simulation relationship between the gyroscope angle and angular rate signals and the motion trajectory when an aircraft is in motion. This makes the system more functionally complete and closer to the actual operation of a complex aircraft system, thereby providing a more realistic gyroscope signal for the test system.

[0003] It should be noted that the information in the background section above is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those skilled in the art. Summary of the Invention

[0004] The purpose of this invention is to provide a method for simulating equivalent signals for gyroscope testing based on the Euler discrete model, thereby overcoming the problem that the gyroscope signal simulation function provided to the testing system is incomplete or the simulation is not realistic enough due to the limitations and defects of related technologies.

[0005] According to one aspect of the present invention, a method for simulating the equivalent signal of a gyroscope test based on an Euler discrete model is provided, comprising the following steps:

[0006] Step S10: First, based on the requirements of the aircraft test mission, generate the aircraft's altitude trajectory data and horizontal trajectory data using a function; or directly use the actual flight altitude trajectory data and horizontal trajectory data; whereby the altitude trajectory data is denoted as... Its representative is The aircraft's altitude trajectory data at any given time, including The time interval for discrete data; horizontal trajectory data is denoted as... Its representative is The horizontal trajectory data of the aircraft at any given moment.

[0007] Step S20: First, the altitude difference signal is calculated based on the altitude trajectory data. Then, the initial state value of the altitude fast differentiator is set, and the altitude difference signal is compared with the altitude difference signal to obtain the altitude differential error signal. This error signal is divided by the time constant of the altitude fast differentiator and superimposed with the altitude nonlinear transformation signal to obtain the state change rate signal of the altitude fast differentiator. This is then integrated to obtain the state signal of the altitude fast differentiator. Next, the horizontal difference signal is calculated based on the horizontal trajectory data. Then, the initial state value of the horizontal fast differentiator is set, and the horizontal difference signal is compared with the horizontal difference signal to obtain the horizontal differential error signal. This error signal is divided by the time constant of the horizontal fast differentiator and superimposed with the horizontal nonlinear transformation signal to obtain the state change rate signal of the horizontal fast differentiator. This is then integrated to obtain the state signal of the horizontal fast differentiator.

[0008] Step S30: Solve for the vertical velocity difference signal based on the state signal of the altitude fast differentiator, then set the initial state value of the vertical velocity fast differentiator, and compare it with the vertical velocity difference signal to obtain the vertical velocity differential error signal. Divide the error signal by the time constant of the vertical velocity fast differentiator and superimpose it with the vertical velocity nonlinear transformation signal to obtain the state change rate signal of the vertical velocity fast differentiator; then integrate to obtain the state signal of the vertical velocity fast differentiator; then solve for the aircraft pitch angular velocity data based on the state signals of the vertical velocity fast differentiator and the horizontal fast differentiator; then integrate to obtain the aircraft pitch angle data.

[0009] Step S40: First, set the initial value of the pitch angle output state of the gyroscope model to 0; then compare it with the initial value of the aircraft pitch angle to obtain the gyroscope measurement error signal; then set the damping ratio parameter and natural frequency parameter of the gyroscope model, design a first-order inertial element, pass the gyroscope measurement error signal through the first-order inertial element to obtain the gyroscope angular rate measurement signal, and then use the Euler method to integrate to obtain the pitch angle output state signal of the gyroscope model.

[0010] Step S50: Based on the gyroscope's drift error constant, input the second-order simulation model of the gyroscope system, obtain the gyroscope drift feedback error signal through feedback, then obtain the gyroscope's constant drift error rate signal through a first-order inertial element, and then perform Euler discrete integration to obtain the gyroscope's constant drift error measurement signal; then, based on the gyroscope's random error signal, input the second-order simulation model of the gyroscope system, obtain the gyroscope's random feedback error signal through feedback, then obtain the gyroscope's random error rate signal through a first-order inertial element, and then perform Euler discrete integration to obtain the gyroscope's random error measurement signal.

[0011] Step S60: Based on the pitch angle output state signal of the gyroscope model, the constant drift error measurement signal of the gyroscope and the random error measurement signal of the gyroscope are superimposed to obtain the final pitch angle equivalent test signal of the aircraft gyroscope system; the angular rate measurement signal of the gyroscope, the constant drift error rate signal of the gyroscope, and the random error rate signal of the gyroscope are superimposed to obtain the final pitch angular velocity equivalent test signal of the aircraft gyroscope system; the final pitch angle equivalent test signal and the final pitch angular velocity equivalent test signal of the aircraft gyroscope system are output to the test system to meet the requirements of the test system.

[0012] In one exemplary embodiment of the present invention, the altitude difference signal is calculated based on altitude trajectory data. Then, the initial state value of the altitude fast differentiator is set, and then compared with the altitude difference signal to obtain the altitude differential error signal. This error signal is divided by the time constant of the altitude fast differentiator and superimposed with the altitude nonlinear transformation signal to obtain the state change rate signal of the altitude fast differentiator. Finally, integration is performed to obtain the state signal of the altitude fast differentiator, including:

[0013] ;

[0014] ;

[0015] ;

[0016] ;

[0017] ;

[0018] in It is a height differential signal. This is a highly differential error signal. It is a highly nonlinear transformed signal. For constant parameters of highly nonlinear transformations; The time constant of a highly fast differentiator; The state change rate signal of a high-speed differentiator; This is the status signal of a high-speed differentiator; The initial state value for the high-speed differentiator can be roughly estimated and set based on the vertical velocity of the aircraft.

[0019] In one exemplary embodiment of the present invention, a horizontal differential signal is calculated based on horizontal trajectory data. Then, an initial state value for the horizontal fast differentiator is set, and this value is compared with the horizontal differential signal to obtain a horizontal differential error signal. This error signal is then divided by the time constant of the horizontal fast differentiator and superimposed with a horizontal nonlinear transformation signal to obtain the state change rate signal of the horizontal fast differentiator. Finally, integration is performed to obtain the state signal of the horizontal fast differentiator, including:

[0020] ;

[0021] ;

[0022] ;

[0023] ;

[0024] ;

[0025] in It is a horizontal differential signal. The horizontal differential error signal, It is a horizontal nonlinear transformation signal. These are constant parameters for the horizontal nonlinear transformation; The time constant of the horizontal fast differentiator; The state change rate signal of the horizontal fast differentiator; This is the status signal of the horizontal fast differentiator; The initial state value for the horizontal fast differentiator can be roughly estimated and set based on the horizontal speed of the aircraft.

[0026] In one exemplary embodiment of the present invention, the vertical velocity difference signal is solved based on the state signal of the altitude fast differentiator. Then, the initial state value of the vertical velocity fast differentiator is set, and then compared with the vertical velocity difference signal to obtain the vertical velocity differential error signal. This error signal is divided by the time constant of the vertical velocity fast differentiator and superimposed with the vertical velocity nonlinear transformation signal to obtain the state change rate signal of the vertical velocity fast differentiator. Integration is then performed to obtain the state signal of the vertical velocity fast differentiator. The aircraft pitch angular velocity data is then solved based on the state signals of the vertical velocity fast differentiator and the horizontal fast differentiator. Integration is then performed to obtain the aircraft pitch angle data, including:

[0027] ;

[0028] ;

[0029] ;

[0030] ;

[0031] ;

[0032] ;

[0033] ;

[0034] in It is a vertical velocity differential signal. This is the vertical velocity differential error signal. This is a nonlinear transformation signal of vertical velocity. These are constant parameters for the nonlinear transformation of the vertical velocity; The time constant of the vertical velocity fast differentiator; The rate of change of state signal of the vertical velocity rapid differentiator; This is the status signal of the vertical velocity rapid differentiator; The initial state value of the vertical velocity rapid differentiator can be roughly estimated and set based on the vertical acceleration of the aircraft; For the aircraft's pitch rate data, For aircraft pitch angle data, is the initial value of the aircraft's pitch angle, and is a constant parameter.

[0035] In one example embodiment of the present invention, the initial value of the pitch angle output state of the gyroscope model is set to 0; then compared with the initial value of the aircraft pitch angle to obtain the gyroscope measurement error signal; then the damping ratio parameter and natural frequency parameter of the gyroscope model are set, a first-order inertial element is designed, and the gyroscope measurement error signal is passed through the first-order inertial element to obtain the gyroscope angular rate measurement rate signal, and then integrated using the Euler method to obtain the gyroscope model pitch angle output state signal, including:

[0036] ;

[0037] ;

[0038] ;

[0039] in This is the measurement error signal for the gyroscope; The damping ratio parameter of the gyroscope model. The natural frequency parameter of the gyroscope model; Let be the transfer function of a first-order inertial element. The differential operator for the transfer function of a first-order inertial element; The gyroscope angular rate measurement rate signal; This is the pitch angle output status signal for the gyroscope model.

[0040] In one exemplary embodiment of the present invention, based on the drift error constant of the gyroscope, a second-order simulation model of the gyroscope system is input to obtain the constant drift error rate signal of the gyroscope, and then Euler discrete integration is performed to obtain the constant drift error measurement signal of the gyroscope; then, based on the random error signal of the gyroscope, a second-order simulation model of the gyroscope system is input to obtain the random error rate signal of the gyroscope, and then Euler discrete integration is performed to obtain the random error measurement signal of the gyroscope, including:

[0041] ;

[0042] ;

[0043] ;

[0044] ;

[0045] ;

[0046] ;

[0047] in This is the drift error constant of the gyroscope. This is the constant drift error rate signal of the gyroscope; This is the constant drift error measurement signal of the gyroscope; This represents the random error signal of the gyroscope. This is the random error rate signal of the gyroscope; This is the random error measurement signal of the gyroscope.

[0048] In one exemplary embodiment of the present invention, outputting the final pitch angle equivalent test signal and the final pitch angular velocity equivalent test signal of the aircraft gyroscope system to the test system includes:

[0049] ;

[0050] ;

[0051] in This is the final equivalent test signal for the pitch angle of the aircraft gyroscope system; This is the equivalent test signal for the pitch angular velocity of the final aircraft gyroscope system.

[0052] Beneficial effects

[0053] This invention presents a method for simulating equivalent signals for gyroscope testing based on the Euler discrete model. Its main innovations are as follows: First, it proposes an approach that establishes the relationship between the aircraft trajectory and the gyroscope output signal, starting from the aircraft trajectory itself. Functionally, this is more detailed and accurate than traditional gyroscope equivalents that only simulate large characteristics such as power-on, power-off, and self-test. Therefore, its approach differs from traditional methods and should more closely approximate the actual state of a real aircraft operating at high altitudes. Second, it establishes a fast differential method for obtaining the gyroscope pitch angle from the aircraft's altitude and horizontal trajectories. This method is more accurate than traditional methods. Particularly noteworthy is that the first-order stage of this method avoids the poor accuracy and abnormally large calculation data in the initial steps of traditional methods, thus avoiding differential spikes in the overall calculation. Traditional methods often exhibit large spikes in the initial calculation steps, significantly reducing simulation accuracy and even rendering the simulation unusable. Thirdly, a discrete model was used to simulate the dynamic measurement characteristics of the gyroscope, including both constant drift error and random error. This method conveniently and realistically simulates the error states of different gyroscopes, allowing for extended use across different gyroscopes simply by modifying the error coefficients according to different systems and gyroscopes. Therefore, this method can be easily extended to different gyroscope systems.

[0054] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Attached Figure Description

[0055] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.

[0056] Figure 1 This is a flowchart of a method for simulating equivalent signals for gyroscope testing based on the Euler discrete model provided by the present invention;

[0057] Figure 2 This is a curve showing the change in altitude trajectory over time (unit: meters) provided by the method in this embodiment of the invention.

[0058] Figure 3 This is a curve showing the change of the horizontal trajectory over time (unit: meters) of the method provided in this embodiment of the invention.

[0059] Figure 4 This is the state signal curve (unitless) of the highly fast differentiator provided in the embodiments of the present invention.

[0060] Figure 5This is the status signal (unit: meters per second) of the horizontal fast differentiator provided in the embodiments of the present invention.

[0061] Figure 6 This is the aircraft pitch angular velocity data curve (unit: degrees per second) provided by the method in the embodiments of the present invention.

[0062] Figure 7 This is the aircraft pitch angle data curve (unit: degrees) of the method provided in the embodiments of the present invention.

[0063] Figure 8 This is the pitch angle output status signal (unit: degrees) of the gyroscope model of the method provided in the embodiments of the present invention.

[0064] Figure 9 This is the constant drift error measurement signal curve of the gyroscope provided by the embodiment of the present invention (unit: degrees).

[0065] Figure 10 This is the random error measurement signal curve (unit: degrees) of the gyroscope provided by the embodiment of the present invention.

[0066] Figure 11 It is the final pitch angular velocity equivalent test signal (unit: degrees per second) of the aircraft gyroscope system provided by the method in the embodiments of the present invention.

[0067] Figure 12 It is the final pitch angle equivalent test signal (unit: degrees) of the aircraft gyroscope system provided in the embodiments of the present invention. Detailed Implementation

[0068] Exemplary embodiments will now be described more fully with reference to the accompanying drawings. However, these exemplary embodiments can be implemented in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided to make the invention more comprehensive and complete, and to fully convey the concept of the exemplary embodiments to those skilled in the art. The described features, structures, or characteristics can be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to give a full understanding of embodiments of the invention. However, those skilled in the art will recognize that the technical solutions of the invention may be practiced with one or more of these specific details omitted, or other methods, components, apparatus, steps, etc., may be employed. In other instances, well-known technical solutions are not shown or described in detail to avoid obscuring various aspects of the invention.

[0069] This invention provides a simulation method for equivalent test signals based on the Euler discrete model. It primarily addresses the problem of generating equivalent gyroscope angle signals for aircraft in the test equivalent generator, enabling high-precision reproduction of the complex dynamic angle changes of an aircraft during flight. This solves the difficulty of accurately simulating gyroscope angle signals during ground testing. First, using predetermined aircraft pitch channel altitude trajectory data, a fast differentiator is used to solve for multiple differentials of the altitude and horizontal trajectories, thereby obtaining the aircraft pitch angle and pitch rate data. Then, the Euler method is used to establish a discrete measurement model of the gyroscope, solving for the measured pitch rate. Considering the effects of random errors and drift errors, integration is performed to obtain the gyroscope's measured pitch angle signal. This outputs high-precision results of the gyroscope angle and angular rate changes during high-altitude flight. This method has the advantages of high simulation accuracy and good agreement with reality, thus meeting the simulation requirements of the testing end for gyroscope signals.

[0070] The following will, with reference to the accompanying drawings, further explain and illustrate the present invention's method for simulating equivalent signals for gyroscope testing based on an Euler discrete model. (Reference) Figure 1 As shown, this method for simulating the equivalent signal of a gyroscope test based on the Euler discrete model may include the following steps:

[0071] Step S10: First, based on the requirements of the aircraft test mission, generate the aircraft's altitude trajectory data and horizontal trajectory data using a function; or directly use the actual flight altitude trajectory data and horizontal trajectory data; whereby the altitude trajectory data is denoted as... Its representative is The aircraft's altitude trajectory data at any given time, including The time interval for discrete data; horizontal trajectory data is denoted as... Its representative is The horizontal trajectory data of the aircraft at any given moment.

[0072] Step S20: First, the altitude difference signal is calculated based on the altitude trajectory data. Then, the initial state value of the altitude fast differentiator is set, and the altitude difference signal is compared with the altitude difference signal to obtain the altitude differential error signal. This error signal is divided by the time constant of the altitude fast differentiator and superimposed with the altitude nonlinear transformation signal to obtain the state change rate signal of the altitude fast differentiator. This is then integrated to obtain the state signal of the altitude fast differentiator. Next, the horizontal difference signal is calculated based on the horizontal trajectory data. Then, the initial state value of the horizontal fast differentiator is set, and the horizontal difference signal is compared with the horizontal difference signal to obtain the horizontal differential error signal. This error signal is divided by the time constant of the horizontal fast differentiator and superimposed with the horizontal nonlinear transformation signal to obtain the state change rate signal of the horizontal fast differentiator. This is then integrated to obtain the state signal of the horizontal fast differentiator.

[0073] Specifically, it can be broken down into the following six steps. The first step is to calculate the altitude difference signal based on the altitude trajectory data, and then compare it with the altitude difference signal to obtain the altitude differential error signal, as follows:

[0074] ;

[0075] ;

[0076] in It is a height differential signal. This is a high-precision differential error signal.

[0077] The second step involves performing a high-level nonlinear transformation on the high-level differential error signal to obtain the high-level nonlinear transformation signal. Then, by dividing the high-level differential error signal by the time constant of the high-speed differentiator and superimposing it with the high-level nonlinear transformation signal, the state change rate signal of the high-speed differentiator is obtained as follows:

[0078] ;

[0079] ;

[0080] in It is a highly nonlinear transformed signal. For constant parameters of highly nonlinear transformations; The time constant of a highly fast differentiator; This is the rate of change signal of the state of a highly fast differentiator.

[0081] The third step involves performing Euler integration on the state change rate signal of the high-fast differentiator to obtain the state signal of the high-fast differentiator as follows:

[0082] ;

[0083] in This is the status signal of a high-speed differentiator; The initial state value for the high-speed differentiator can be roughly estimated and set based on the vertical velocity of the aircraft.

[0084] The fourth step is to solve for the horizontal difference signal based on the horizontal trajectory data, then set the initial state value of the horizontal fast differentiator, and compare it with the horizontal difference signal to obtain the horizontal differential error signal as follows:

[0085] ;

[0086] ;

[0087] in It is a horizontal differential signal. This is the horizontal differential error signal.

[0088] Fifth, perform a horizontal nonlinear transformation on the horizontal differential error signal to obtain the horizontal nonlinear transformation signal; then, divide the horizontal differential error signal by the time constant of the horizontal fast differentiator and superimpose the horizontal nonlinear transformation signal to obtain the state change rate signal of the horizontal fast differentiator as follows:

[0089] ;

[0090] ;

[0091] in It is a horizontal nonlinear transformation signal. These are constant parameters for the horizontal nonlinear transformation; The time constant of the horizontal fast differentiator; This is the rate of change signal of the horizontal fast differentiator.

[0092] Step 6: Integrate the state change rate signal of the horizontal fast differentiator to obtain the state signal of the horizontal fast differentiator as follows:

[0093] ;

[0094] in This is the status signal of the horizontal fast differentiator; The initial state value for the horizontal fast differentiator can be roughly estimated and set based on the horizontal speed of the aircraft.

[0095] Step S30: Solve for the vertical velocity difference signal based on the state signal of the altitude fast differentiator, then set the initial state value of the vertical velocity fast differentiator, and compare it with the vertical velocity difference signal to obtain the vertical velocity differential error signal. Divide the error signal by the time constant of the vertical velocity fast differentiator and superimpose it with the vertical velocity nonlinear transformation signal to obtain the state change rate signal of the vertical velocity fast differentiator; then integrate to obtain the state signal of the vertical velocity fast differentiator; then solve for the aircraft pitch angular velocity data based on the state signals of the vertical velocity fast differentiator and the horizontal fast differentiator; then integrate to obtain the aircraft pitch angle data.

[0096] Specifically, it can be broken down into the following five steps. The first step is to solve for the vertical velocity difference signal based on the state signal of the height-speed fast differentiator, then set the initial state value of the vertical velocity fast differentiator, and finally compare it with the vertical velocity difference signal to obtain the vertical velocity differential error signal as follows:

[0097] ;

[0098] ;

[0099] in It is a vertical velocity differential signal. This is the vertical velocity differential error signal.

[0100] The second step involves performing a horizontal nonlinear transformation on the vertical velocity differential error signal to obtain the vertical velocity nonlinear transformation signal. Then, by dividing the vertical velocity differential error signal by the time constant of the vertical velocity fast differentiator and superimposing it with the vertical velocity nonlinear transformation signal, the state change rate signal of the vertical velocity fast differentiator is obtained as follows:

[0101] ;

[0102] ;

[0103] in This is a nonlinear transformation signal of vertical velocity. These are constant parameters for the nonlinear transformation of the vertical velocity; The time constant of the vertical velocity fast differentiator; This is the rate of change signal of the vertical velocity fast differentiator.

[0104] The third step involves integrating the state change rate signal of the vertical velocity rapid differentiator to obtain the state signal of the vertical velocity rapid differentiator as follows:

[0105] ;

[0106] in This is the status signal of the vertical velocity rapid differentiator; The initial value for the vertical velocity rapid differentiator can be roughly estimated and set based on the vertical acceleration of the aircraft.

[0107] The fourth step involves calculating the aircraft's pitch angular velocity data based on the state signals of the vertical and horizontal rapid differentiators, as follows:

[0108] ;

[0109] in This is the pitch rate data for the aircraft.

[0110] Fifth, integrate the aircraft's pitch angular velocity data to obtain the aircraft's pitch angle data as follows:

[0111] ;

[0112] in For aircraft pitch angle data, is the initial value of the aircraft's pitch angle, and is a constant parameter.

[0113] Step S40: First, set the initial value of the pitch angle output state of the gyroscope model to 0; then compare it with the initial value of the aircraft pitch angle to obtain the gyroscope measurement error signal; then set the damping ratio parameter and natural frequency parameter of the gyroscope model, design a first-order inertial element, pass the gyroscope measurement error signal through the first-order inertial element to obtain the gyroscope pitch angle rate measurement signal, and then use the Euler method to integrate to obtain the gyroscope model pitch angle output state signal.

[0114] Specifically, this can be broken down into the following three steps. First, set the initial value of the gyroscope model's pitch angle output state to 0; then compare it with the initial value of the aircraft's pitch angle to obtain the gyroscope measurement error signal as follows:

[0115] ;

[0116] in This is the measurement error signal for the gyroscope.

[0117] The second step is to set the damping ratio and natural frequency parameters of the gyroscope model, design a first-order inertial element, and pass the gyroscope measurement error signal through the first-order inertial element to obtain the gyroscope angular rate measurement signal as follows:

[0118] ;

[0119] in The damping ratio parameter of the gyroscope model. The natural frequency parameter of the gyroscope model; Let be the transfer function of a first-order inertial element. The differential operator for the transfer function of a first-order inertial element.

[0120] The third step is to use the Euler method to integrate the gyroscope pitch rate measurement signal to obtain the gyroscope model pitch angle output state signal.

[0121] ;

[0122] in The gyroscope angular rate measurement rate signal; This is the pitch angle output status signal for the gyroscope model.

[0123] Step S50: Based on the gyroscope's drift error constant, input the second-order simulation model of the gyroscope system, obtain the gyroscope drift feedback error signal through feedback, then obtain the gyroscope's constant drift error rate signal through a first-order inertial element, and then perform Euler discrete integration to obtain the gyroscope's constant drift error measurement signal; then, based on the gyroscope's random error signal, input the second-order simulation model of the gyroscope system, obtain the gyroscope's random feedback error signal through feedback, then obtain the gyroscope's random error rate signal through a first-order inertial element, and then perform Euler discrete integration to obtain the gyroscope's random error measurement signal.

[0124] Specifically, this can be broken down into the following six steps. The first step involves comparing the gyroscope's drift error constant with the constant drift error measurement signal obtained from the second-order simulation model of the input gyroscope system. The resulting gyroscope drift feedback error signal is as follows:

[0125] ;

[0126] in This is the drift error constant of the gyroscope. This is the gyroscope drift feedback error signal.

[0127] The second step is to pass the gyroscope drift error signal through a first-order inertial element to obtain the gyroscope's constant drift error rate signal as follows:

[0128] ;

[0129] in This is the constant drift error rate signal of the gyroscope.

[0130] The third step involves performing Euler discrete integration on the constant drift error rate signal of the gyroscope to obtain the constant drift error measurement signal of the gyroscope, as follows:

[0131] ;

[0132] in This is the constant drift error measurement signal for the gyroscope.

[0133] The fourth step involves comparing the random error signal of the gyroscope with the random error measurement signal of the gyroscope obtained after inputting the second-order simulation model of the gyroscope system. The resulting random feedback error signal of the gyroscope is as follows:

[0134] ;

[0135] in This represents the random error signal of the gyroscope. This is the random feedback error signal of the gyroscope.

[0136] The fifth step involves passing the gyroscope's random feedback error signal through the following first-order inertial element to obtain the gyroscope's random error rate signal:

[0137] ;

[0138] in This is the random error rate signal of the gyroscope.

[0139] The sixth step is to integrate the random error rate signal of the gyroscope to obtain the random error measurement signal of the gyroscope.

[0140] ;

[0141] in This is the random error measurement signal of the gyroscope.

[0142] Step S60: Based on the pitch angle output state signal of the gyroscope model, the constant drift error measurement signal of the gyroscope and the random error measurement signal of the gyroscope are superimposed to obtain the final pitch angle equivalent test signal of the aircraft gyroscope system; the angular rate measurement signal of the gyroscope, the constant drift error rate signal of the gyroscope, and the random error rate signal of the gyroscope are superimposed to obtain the final pitch angular velocity equivalent test signal of the aircraft gyroscope system; the final pitch angle equivalent test signal and the final pitch angular velocity equivalent test signal of the aircraft gyroscope system are output to the test system, thus satisfying the requirements of the test system as follows:

[0143] ;

[0144] ;

[0145] in This is the final equivalent test signal for the pitch angle of the aircraft gyroscope system; This is the equivalent test signal for the pitch angular velocity of the final aircraft gyroscope system.

[0146] Case Implementation and Computer Simulation Results Analysis

[0147] In step S10, the variation law of the altitude trajectory over time is set as follows: The horizontal speed of the aircraft is Therefore, its horizontal trajectory is The altitude trajectory changing over time is shown in the curve. Figure 2 As shown, the curve of the horizontal trajectory changing over time is as follows: Figure 3 As shown.

[0148] In step S20, select , , , The state signal of the high-speed differentiator is obtained as follows: Figure 4 As shown; the state signal of the horizontal fast differentiator is as follows Figure 5 As shown.

[0149] In step S30, select , The aircraft pitch angular velocity data is obtained as follows: Figure 6 As shown, the aircraft pitch angle data is as follows: Figure 7 As shown.

[0150] In step S40, select , The pitch angle output state signal of the gyroscope model is obtained as follows: Figure 8 As shown.

[0151] In step S50, select The constant drift error measurement signal of the gyroscope is obtained as follows: Figure 9 As shown. In step S60, select The random signal is Gaussian distributed; the random error measurement signal of the gyroscope is obtained as follows: Figure 10 As shown.

[0152] In step S60, the final equivalent test signal of the pitch angular velocity of the aircraft gyroscope system is obtained, as shown below. Figure 11 As shown; the final pitch angle equivalent test signal of the aircraft gyroscope system is obtained as follows: Figure 12 As shown.

[0153] Depend on Figure 7 It can be seen that traditional methods often exhibit significant spikes in the initial few steps of the solution process, with values ​​reaching thousands or tens of thousands, thus affecting the accuracy of the initial solution; however, they quickly return to normal afterward. In contrast,... Figure 7 Zhongyu Figure 11 as well as Figure 12 No large spikes were found in the angle data, indicating that the method provided by this invention has high engineering application value.

Claims

1. A method for simulating the equivalent signal of a gyroscope test based on the Euler discrete model, characterized by the following steps: Step S10, first, according to the aircraft test task background needs, using function to generate the height trajectory data and horizontal trajectory data of the aircraft; or directly using the height trajectory data and horizontal trajectory data of the real flight; wherein the height trajectory data is denoted as , which represents the height trajectory data of the aircraft at the moment of , wherein is the time interval of discrete data; the horizontal trajectory data is denoted as , which represents the horizontal trajectory data of the aircraft at the moment of ; Step S20: First, the altitude difference signal is calculated based on the altitude trajectory data. Then, the initial state value of the altitude fast differentiator is set, and the altitude difference signal is compared with the altitude difference signal to obtain the altitude differential error signal. This error signal is divided by the time constant of the altitude fast differentiator and superimposed with the altitude nonlinear transformation signal to obtain the state change rate signal of the altitude fast differentiator. Integration is then performed to obtain the state signal of the altitude fast differentiator. Next, the horizontal difference signal is calculated based on the horizontal trajectory data. Then, the initial state value of the horizontal fast differentiator is set, and the horizontal difference signal is compared with the horizontal difference signal to obtain the horizontal differential error signal. This error signal is divided by the time constant of the horizontal fast differentiator and superimposed with the horizontal nonlinear transformation signal to obtain the state change rate signal of the horizontal fast differentiator. Integration is then performed to obtain the state signal of the horizontal fast differentiator. ； ； ； ； ； ； ； ； ； ； in It is a height differential signal. This is a highly differential error signal. It is a highly nonlinear transformed signal. For highly nonlinear transformations, constant parameters are used. The time constant of a highly fast differentiator; The state change rate signal of a high-speed differentiator; This is the status signal of a high-speed differentiator; The initial state value of the high-speed differentiator can be roughly estimated and set based on the vertical velocity of the aircraft; It is a horizontal differential signal. The horizontal differential error signal, It is a horizontal nonlinear transformation signal. These are constant parameters for the horizontal nonlinear transformation; The time constant of the horizontal fast differentiator; The state change rate signal of the horizontal fast differentiator; This is the status signal of the horizontal fast differentiator; The initial state value of the horizontal fast differentiator can be roughly estimated and set based on the horizontal speed of the aircraft. Step S30: Solve the vertical velocity difference signal based on the state signal of the altitude fast differentiator, then set the initial state value of the vertical velocity fast differentiator, and compare it with the vertical velocity difference signal to obtain the vertical velocity differential error signal. Divide the error signal by the time constant of the vertical velocity fast differentiator and superimpose the vertical velocity nonlinear transformation signal to obtain the state change rate signal of the vertical velocity fast differentiator; then integrate to obtain the state signal of the vertical velocity fast differentiator; finally, solve the aircraft pitch angular velocity data based on the state signals of the vertical velocity fast differentiator and the horizontal fast differentiator. Then, integration is performed to obtain the aircraft's pitch angle data; ； ； ； ； ； ； ； in It is a vertical velocity differential signal. This is the vertical velocity differential error signal. This is a nonlinear transformation signal of vertical velocity. These are constant parameters for the nonlinear transformation of the vertical velocity; The time constant of the vertical velocity fast differentiator; The rate of change of state signal of the vertical velocity fast differentiator; This is the status signal of the vertical velocity rapid differentiator; The initial state value of the vertical velocity rapid differentiator can be roughly estimated and set based on the vertical acceleration of the aircraft; For the aircraft's pitch rate data, For aircraft pitch angle data, is the initial value of the aircraft's pitch angle, and is a constant parameter; Step S40: First, set the initial value of the pitch angle output state of the gyroscope model to 0; then compare it with the initial value of the aircraft pitch angle to obtain the gyroscope measurement error signal. Next, the damping ratio and natural frequency parameters of the gyroscope model are set, and a first-order inertial element is designed. The gyroscope measurement error signal is passed through the first-order inertial element to obtain the gyroscope angular rate measurement signal. Then, the Euler method is used for integration to obtain the pitch angle output state signal of the gyroscope model as follows: ； ； ； in This is the measurement error signal for the gyroscope; The damping ratio parameter of the gyroscope model. The natural frequency parameter of the gyroscope model; Let be the transfer function of a first-order inertial element. The differential operator for the transfer function of a first-order inertial element; This is the angular rate measurement signal from the gyroscope; The pitch angle output status signal of the gyroscope model; Step S50: Based on the gyroscope's drift error constant, input the second-order simulation model of the gyroscope system, obtain the gyroscope drift feedback error signal through feedback, then obtain the gyroscope's constant drift error rate signal through a first-order inertial element, and then perform Euler discrete integration to obtain the gyroscope's constant drift error measurement signal; then, based on the gyroscope's random error signal, input the second-order simulation model of the gyroscope system, obtain the gyroscope's random feedback error signal through feedback, then obtain the gyroscope's random error rate signal through a first-order inertial element, and then perform Euler discrete integration to obtain the gyroscope's random error measurement signal; ； ； ； ； ； ； in This is the drift error constant of the gyroscope. This is the gyroscope drift feedback error signal. This is the random feedback error signal from the gyroscope. This is the constant drift error rate signal of the gyroscope; This is the constant drift error measurement signal of the gyroscope; This represents the random error signal of the gyroscope. This is the random error rate signal of the gyroscope; This is the random error measurement signal of the gyroscope; Step S60: Based on the pitch angle output state signal of the gyroscope model, the constant drift error measurement signal of the gyroscope and the random error measurement signal of the gyroscope are superimposed to obtain the final pitch angle equivalent test signal of the aircraft gyroscope system; the angular rate measurement signal of the gyroscope, the constant drift error rate signal of the gyroscope, and the random error rate signal of the gyroscope are superimposed to obtain the final pitch angular velocity equivalent test signal of the aircraft gyroscope system; the final pitch angle equivalent test signal and the final pitch angular velocity equivalent test signal of the aircraft gyroscope system are output to the test system, thus satisfying the requirements of the test system as follows: ； ； in This is the final equivalent test signal for the pitch angle of the aircraft gyroscope system; This is the final equivalent test signal for the pitch angular velocity of the aircraft gyroscope system.

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