A measurement platform and identification method for aircraft moment of inertia

Abstract

The invention discloses an aircraft rotational inertia measurement platform and a method for identifying measurement data, comprising an aircraft mounting plate, a connecting flange, a first support plate, a first transmission optical axis, a first plum blossom coupling, a dynamic torque sensor, a second support plate, a support frame, a second plum blossom coupling, a third support plate, a second transmission optical axis, a third plum blossom coupling, a fourth support plate, an Omron encoder, and a fixing frame; the overall structural material of the invention is an aluminum profile, which is stable and reliable; the mounting plate is a carbon fiber plate, which is light in weight, can withstand high pressure, is not easily deformed, and is convenient for aircraft installation and testing.

Description

Technical Field

[0001] The invention belongs to the technical field of moment of inertia measurement, and in particular relates to an aircraft moment of inertia measurement platform and an identification method. Background Art

[0002] The moment of inertia is a measure of the inertia of a rigid body when it rotates around an axis. It can be used to establish the relationship between angular momentum, angular velocity, torque, and angular acceleration. Its value depends on the shape of the object, mass distribution, and the position of the axis of rotation. The moment of inertia of a uniform rigid body with a regular shape can be directly calculated using a formula, but for the moment of inertia of an irregular or non-uniform rigid body, it is difficult to calculate an accurate result with a theoretical formula, and it is generally measured by experimental methods.

[0003] In the aircraft system, the moment of inertia can directly reflect the current heading information of the aircraft, which is crucial for attitude control and state estimation. The data measured by the moment of inertia measurement platform is directly applied to the flight parameter adjustment after identification and processing, which can effectively promote the safe and stable operation of the aircraft system. Traditional moment of inertia measurement methods mainly include three-wire pendulum method, torsion pendulum method and compound pendulum method, etc., which are not effective for aircraft moment of inertia testing, especially for the linear aircraft targeted by the present invention. The above methods are not easy to operate, the measured data has large errors, and it is not easy to directly identify and apply to the flight control microcontroller end. Summary of the invention

[0004] Based on the above-mentioned deficiencies of the existing measurement technology, the present invention discloses an aircraft rotational inertia measurement platform and identification method which are flexible in operation, high in accuracy and convenient in application of identification data.

[0005] The technical scheme of the present invention includes: an aircraft rotational inertia measurement platform, including an aircraft mounting plate 1, a connecting flange 2, a first support plate 3, a first transmission optical axis 4, a first plum blossom coupling 5, a dynamic torque sensor 6, a second support plate 7, a support frame 8, a second plum blossom coupling 9, a third support plate 10, a second transmission optical axis 11, a third plum blossom coupling 12, a fourth support plate 13, an Omron encoder 14, and a fixing frame 15; the aircraft mounting plate 1 is connected to the first transmission optical axis 4 through the connecting flange 2, and there is a small gap between the aircraft mounting plate 1 and the support frame 8, so that the aircraft can rotate without damping on the aircraft mounting plate 1; the first transmission optical axis 4 is connected to the dynamic torque sensor 6 through the first plum blossom coupling 5; the dynamic torque sensor 6 is respectively connected to the second plum blossom coupling 9, The second transmission optical axis 11 and the third plum blossom coupling 12 are connected to the Omron encoder 14; the two ends of the horizontally arranged first support plate 3, the second support plate 7, the third support plate 10 and the fourth support plate 13 are connected to the vertical support frame 8 in sequence from top to bottom in spatial order, and the support frame 8 is connected to the fixed frame 15 placed horizontally on the ground through an angle code, and the overall structure is vertically facing the ground; wherein the aircraft mounting plate 1 and the connecting flange 2 are located above the first support plate 3; the first transmission optical axis 4, the first plum blossom coupling 5 and the dynamic torque sensor 6 are located between the first support plate 3 and the second support plate 7; the second plum blossom coupling 9 is located between the second support plate 7 and the third support plate 10; the third plum blossom coupling 12 is located between the third support plate 10 and the fourth support plate 13.

[0006] Furthermore, corresponding mounting holes are provided on the aircraft mounting plate 1 according to the requirements of the aircraft multi-axis test, the aircraft is fixed by a combination of fasteners and pipe clamps, and the fasteners are used to hard connect with the connecting flange 2 .

[0007] Furthermore, the first plum blossom coupling 5 , the second plum blossom coupling 9 , and the third plum blossom coupling 12 are divided into two sizes, respectively adapting to different shaft diameters of the first transmission optical axis 4 , the dynamic torque sensor 6 , the second transmission optical axis 11 , and the Omron encoder 14 .

[0008] Furthermore, the first transmission optical axis 4 and the second transmission optical axis 11 are both limited by a diamond-shaped optical axis fixing seat, and the fixing seat is rigidly connected to the first support plate 3 , the second support plate 7 , the third support plate 10 , and the fourth support plate 13 using fasteners.

[0009] Further, the dynamic torque sensor 6 is mounted on the second support plate 7 , and the Omron encoder 14 is mounted on the fourth support plate 13 .

[0010] A method for identifying measurement data of an aircraft moment of inertia measurement platform of the present invention comprises the following steps:

[0011] Step 1, fix the aircraft to be tested on the aircraft mounting plate 1, check and correct the transition gap between the aircraft mounting plate 1 and the support frame 8, and rotate the aircraft to perform multi-axis motion;

[0012] Step 2, the dynamic torque sensor 6 displays information such as torque, power and speed generated by the rotation of the aircraft in step 1, and transmits it to the data processing unit through a signal line;

[0013] Step 3, the Omron encoder 14 converts the rotation information into a pulse signal, counts the rising edge (falling edge) pulse signal accumulation value of the pulse signal, calculates the corresponding rotation angular velocity by proportion, and obtains the angular acceleration by differentiating the tracking differentiator;

[0014] Step 4, using the Kalman filter parameter identification method, establish the state equation for the torque and angular acceleration measured in steps 2 and 3, use a mathematical model to describe the relationship between the aircraft's moment of inertia and the torque and angular acceleration, and solve the identification parameters.

[0015] The present invention also proposes a method for identifying aircraft moment of inertia measurement data, which is a parameter identification method based on Kalman filtering. For aircraft systems with large measured noise and random interference and accompanied by variable parameters, the Kalman filter is an optimal random state estimator with strong anti-interference and noise smoothing capabilities, and has a good effect on the estimation of system states and time-varying parameters. The Kalman filter parameter identification algorithm is based on the establishment of a random system state space model, and performs optimal estimation and measurement calculation on the system state vector, measurement vector, and filter gain to obtain the corresponding system identification state equation, and uses a mathematical model to describe the relationship between the aircraft moment of inertia and torque and rotational angular acceleration.

[0016] The present invention has the following technical effects:

[0017] (1) In view of the wide variety of aircraft types, the invention is not only applicable to traditional multi-rotor aircraft such as quad-rotors and hexacopters, but also to "one" linear multi-rotor aircraft.

[0018] (2) In addition to aircraft, the platform of the present invention can measure the moment of inertia of a variety of rigid bodies within the load range, and has outstanding characteristics and universal applicability.

[0019] (3) The overall structural material is aluminum profile, which is stable and reliable; the mounting plate is made of carbon fiber plate, which is lightweight and can withstand high pressure, is not easily deformed, and is convenient for aircraft installation and testing.

[0020] (4) The system parameter identification method is reliable, especially for the complexity and diversity of aircraft systems, and can quickly process information and establish corresponding mathematical models. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 This is the overall structure diagram of the aircraft moment of inertia measurement platform;

[0022] Figure 2 This is a detail view of the support plate;

[0023] Figure 3 This is a detailed picture of the dynamic torque sensor;

[0024] Figure 4 This is the flow chart of Kalman filter parameter identification.

[0025] exist Figure 1 In the figure, 1-aircraft mounting plate; 2-connecting flange; 3-first support plate; 4-first transmission optical axis; 5-first plum blossom coupling; 6-dynamic torque sensor; 7-second support plate; 8-support frame; 9-second plum blossom coupling; 10-third support plate; 11-second transmission optical axis; 12-third plum blossom coupling; 13-fourth support plate; 14-Omron encoder; 15-fixed frame. DETAILED DESCRIPTION

[0026] The technical solutions in the embodiments of the present invention will be described clearly and completely below in conjunction with the accompanying drawings in the embodiments of the present invention.

[0027] like Figure 1The overall structure diagram of the aircraft rotational inertia measurement platform shown in the figure is mainly composed of the following parts: 1-aircraft mounting plate; 2-connecting flange; 3-first support plate; 4-first transmission optical axis; 5-first plum blossom coupling; 6-dynamic torque sensor; 7-second support plate; 8-support frame; 9-second plum blossom coupling; 10-third support plate; 11-second transmission optical axis; 12-third plum blossom coupling; 13-fourth support plate; 14-Omron encoder; 15-fixed frame. In the whole structure, there are multiple mounting holes on the aircraft mounting plate 1. In addition to the fixing holes for the aircraft multi-axis test, fixing holes for other objects to be tested are reserved, which has functional versatility; the aircraft mounting plate 1 is connected to the first transmission optical axis 4 through the connecting flange 2, and the first plum blossom coupling 5 connects the first transmission optical axis 4 and the upper rotating axis of the dynamic torque sensor 6; in particular, there is a gap between the aircraft mounting plate 1 and the support frame 8, and there is no direct contact, so as to ensure that no additional friction is generated when the aircraft rotates, avoiding errors in hardware connection; the first transmission optical axis 4 and the first plum blossom coupling The device 5 transmits the rotation data losslessly to the dynamic torque sensor 6. The dynamic torque sensor 6 has a digital display function, which directly displays the torque, power, speed and other information generated by the current rotation of the aircraft; the lower rotating shaft of the dynamic torque sensor 6 is connected to the Omron encoder 14 through the second plum blossom coupling 9, the second transmission optical axis 11 and the third plum blossom coupling 12, and the aircraft rotation data is fed back to the rotating shaft of the Omron encoder 14; the Omron encoder 14 is an incremental encoder, which can output information in the form of pulses, which are pulled up to the single-chip microcomputer module for identification and processing.

[0028] like Figure 2 The main function of the first support plate 3, the second support plate 7, the third support plate 10, and the fourth support plate 13 shown is to connect the support frame 8 to reinforce the moment of inertia measurement platform of the present invention. At the same time, diamond-shaped optical axis fixing seats are installed on the upper and lower surfaces to limit and reinforce the first transmission optical axis 4 and the second transmission optical axis 11; the support frame 8 is an aluminum profile perpendicular to the ground, and is connected to the fixing frame 15 through an angle code to stabilize the platform. At the same time, the fixing frame 15 is a long-axis aluminum profile, and heavy objects can be placed on the surface for reinforcement, or the platform of the present invention can be fixed to the horizontal ground through a riding card.

[0029] like Figure 3 Shown is a detailed view of the dynamic torque sensor 6. The sensor has a digital display function, which directly displays the current torque, power, speed and other information. The above information can be exported through the built-in aviation plug; it has two upper and lower rotating axes, which can transmit the rotation data losslessly to the lower end connector.

[0030] The upper end of the Omron encoder 14 has a rotating shaft, and a multi-phase output rotation angle pulse signal. The forward and reverse information is obtained by judging the leading and lagging conditions of the A and B phases. At the same time, the rising edge (falling edge) pulse signal accumulation value is counted, and the corresponding rotation angular velocity is obtained after proportional calculation. The angular acceleration is obtained by differentiating the tracking differentiator. The details are as follows:

[0031] Assume the system discrete form is:

[0032]

[0033] Where u is the input signal; r is the adjustable input range; x1(k) and x2(k) are the output signals; and h is the integration step.

[0034] Solving the “fast discrete tracking differentiator” for equation (1) yields:

[0035]

[0036] Where t is time.

[0037] u=f(x1,x2,r,h)

[0038] d=rh,d0=dh

[0039]

[0040]

[0041]

[0042] Where y is a parameter variable. The variable h in the function f(·) is replaced by a new variable h0 that is independent of the step size h, and we get

[0043]

[0044] Wherein r and h0 are adjustable parameters, v(t) is an input variable, corresponding to the rotational angular velocity information after the output processing of the Omron encoder 14.

[0045] like Figure 4 The figure shows the Kalman filter parameter identification flow chart. The Kalman filter parameter identification algorithm is based on the establishment of a random system state space model, and performs optimal estimation and measurement calculations on the system state vector, measurement vector, and filter gain to obtain the corresponding system identification state equation, and uses a mathematical model to describe the relationship between the aircraft's moment of inertia, torque, and angular acceleration. The details are as follows:

[0046] Construct a state-space model of a random system:

[0047] X k =Φk / k-1 X k-1 +Γ k / k-1 W k-1 (4)

[0048] Z k =H k X k +V k (5)

[0049] Where, X k is the state vector; Z k is the measurement vector; Φ k / k-1 is the state one-step transfer matrix; Γ k / k-1 is the system noise allocation matrix; H k is the measurement matrix. k-1 、V k They are the system noise vector at time k-1 and the measurement noise vector at time k, both of which are zero-mean Gaussian white noise vector sequences (obeying normal distribution) and are uncorrelated, that is, they satisfy

[0050]

[0051] Where W k is the system noise vector at time k; E is the expectation; Q k is the system noise covariance matrix; R k is the measurement noise covariance matrix; δ kj represents the Dirac impulse function, specifically:

[0052]

[0053] Formula (6) requires Q k is non-negative definite and R k is positive definite, that is, Q k ≥0 and R k >0.

[0054] The steps of the linear Kalman filter algorithm are as follows:

[0055] (1) One-step state prediction

[0056]

[0057] (2) Corrected state prediction value

[0058]

[0059] (3) Calculate filter gain

[0060]

[0061] (4) State one-step prediction mean square error matrix

[0062]

[0063] (5) Correct the predicted state mean square error matrix

[0064] P k =(IK k H k ) k / k-1 (12)

[0065] In the formula, is the optimal estimate of the state at time k-1; is the best one-step prediction; K k is the filter gain; P k for The mean square error matrix.

[0066] The aircraft of the present invention is mechanically fixedly connected to the mounting plate, and the relative displacement between the two is extremely small and can be ignored when analyzing the rotation equation. The rotation equation is specifically:

[0067]

[0068] Where, J is the moment of inertia; ω r is the angular velocity; α is the angular acceleration; M is the torque.

[0069] Wherein, M is the value read by the dynamic torque sensor 6, ω r It is obtained by the Omron encoder 14 after being processed by the above specific implementation method, and α is obtained after being differentiated by the tracking differentiator.

[0070] Establish the state space model of the aircraft moment of inertia measurement platform:

[0071]

[0072] The optimal covariance estimate is solved according to the above-mentioned linear Kalman filter algorithm steps to complete the parameter identification of the present invention.

[0073] In the description of this specification, the description with reference to the terms "one embodiment", "some embodiments", "illustrative embodiments", "examples", "specific examples", or "some examples" means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representation of the above terms does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described may be combined in any one or more embodiments or examples in a suitable manner.

[0074] Although the embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and variations may be made to the embodiments without departing from the principles and spirit of the present invention, and that the scope of the present invention is defined by the claims and their equivalents.

Claims

1. A measurement data identification method for an aircraft moment of inertia measurement platform, characterized in that: The platform comprises an aircraft mounting plate (1), a connecting flange (2), a first supporting plate (3), a first transmission optical axis (4), a first plum blossom coupling (5), a dynamic torque sensor (6), a second supporting plate (7), a supporting frame (8), a second plum blossom coupling (9), a third supporting plate (10), a second transmission optical axis (11), a third plum blossom coupling (12), a fourth supporting plate (13), an Omron encoder (14), and a fixing frame (15); the aircraft mounting plate (1) is connected to the first transmission optical axis (4) through the connecting flange (2), and a small gap is formed between the first transmission optical axis (4) and the supporting frame (8), so that the aircraft can rotate without damping on the aircraft mounting plate (1); the first transmission optical axis (4) is connected to the dynamic torque sensor (6) through the first plum blossom coupling (5); the dynamic torque sensor (6) is connected to the Omron encoder (14) through the second plum blossom coupling (9), the second transmission optical axis (11), and the third plum blossom coupling (12) respectively. The first support plate (3), the second support plate (7), the third support plate (10) and the fourth support plate (13) are connected at both ends in a spatial order from top to bottom to a vertical support frame (8), and the support frame (8) is connected to a fixing frame (15) placed horizontally on the ground through an angle code, and the overall structure is vertically facing the ground; wherein the aircraft mounting plate (1) and the connecting flange (2) are located above the first support plate (3); the first transmission optical axis (4), the first plum blossom coupling (5) and the dynamic torque sensor (6) are located between the first support plate (3) and the second support plate (7); the second plum blossom coupling (9) is located between the second support plate (7) and the third support plate (10); the third plum blossom coupling (12) is located between the third support plate (10) and the fourth support plate (13); the dynamic torque sensor (6) is installed on the second support plate (7), and the Omron encoder (14) is installed on the fourth support plate (13); The method comprises the following steps: Step 1, fix the aircraft to be tested on the aircraft mounting plate (1), check and correct the transition gap between the aircraft mounting plate (1) and the support frame (8), and rotate the aircraft to perform multi-axis motion; Step 2, the dynamic torque sensor (6) displays information such as torque, power and speed generated by the rotation of the aircraft in step 1, and transmits the information to the data processing unit through a signal line; Step 3, the Omron encoder (14) converts the rotation information into a pulse signal, counts the rising edge or falling edge pulse signal accumulation value of the pulse signal, calculates the corresponding rotation angular velocity by proportion, and obtains the angular acceleration by differentiating with a tracking differentiator; Step 4, using the Kalman filter parameter identification method, establish the state equation for the torque and angular acceleration measured in steps 2 and 3, use a mathematical model to describe the relationship between the aircraft's moment of inertia and the torque and angular acceleration, and solve the identification parameters.

2. The method according to claim 1, characterized in that Corresponding mounting holes are arranged on the aircraft mounting plate (1) according to the requirements of the aircraft multi-axis test, the aircraft is fixed by a combination of fasteners and pipe clamps, and the fasteners are hard-connected to the connecting flange (2).

3. The method according to claim 1, characterized in that The first plum blossom coupling (5), the second plum blossom coupling (9) and the third plum blossom coupling (12) are divided into two sizes, respectively adapted to the different shaft diameters of the first transmission optical axis (4), the dynamic torque sensor (6), the second transmission optical axis (11) and the Omron encoder (14).

4. The method according to claim 1, characterized in that: The first transmission optical axis (4) and the second transmission optical axis (11) are both limited by a rhombus-shaped optical axis fixing seat, and the fixing seat is rigidly connected to the first support plate (3), the second support plate (7), the third support plate (10), and the fourth support plate (13) by using fasteners.

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